Airbus A220 Technical Training Manual - Powerplant Bombardier Cseries Cs300
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BD500 SERIES (PW PW1500G) INITIAL MAINTENANCE TRAINING COURSE
TECHNICAL TRAINING MANUAL MODULE 5: POWER PLANT CMUI: CS130-21.07-06-00-121418 VERSION V6.00 BD-500-1A10 BD-500-1A11
CUSTOMER TRAINING Montreal Training Centre 8575 Côte-de-Liesse Road Saint-Laurent, Québec, Canada H4T 1G5 Telephone (514) 344-6620 Toll-Free North America 1 (877) 551-1550 Fax (514) 344-6643 www.batraining.com iflybombardier.com
Reader Notice and Disclaimer Please note that this version of C Series Module 5: Power Plant is up-to-date as of the date of the training course(s) that you are attending, and may only be used for the purpose of such course(s). Bombardier disclaims responsibility in case of use of the document for any other purpose than said course(s). If you would like to continue using C Series Module 5: Power Plant after your subscription access has expired, you are required to register and purchase a subscription plan in order to receive the necessary updates when they become available at the following website: http://www.batraining.com. Bombardier Inc., or its subsidiaries (collectively “Bombardier”), provides this information to its customers and to government authorities in confidence. The information contained herein must therefore be treated as proprietary confidential information, and as such it must be excluded from any request for access to a record pursuant to section 20 of the Access to Information Act, RSC 1985, c A-1, or any other applicable statute with respect to access to information. Public release of this information would be highly detrimental to Bombardier and as such is strictly prohibited without Bombardier's prior written authorization. This document, which comprises protected intellectual property and trade secrets, shall not be used, reproduced, published, broadcasted, copied, translated, distributed, transferred, stored on any medium, including in a retrieval system, communicated, altered, or converted in any form or by any means, electronic or otherwise, in whole or in part, without Bombardier's prior written authorization. The rights to all patents, inventions, know-how, copyrights, trademarks, trade secrets, registered designs, database rights, semiconductor topography rights, service marks, logos, domain names, business names, trade names, moral rights and all related registrations or applications in any country or jurisdiction contained herein belong to or are used under license by Bombardier. This documentation, the technical data it contains, and all other information shall not be modified, translated, reverse assembled, reverse engineered, or decompiled and shall be used solely for training purposes. Nothing contained herein shall be constructed as granting, explicitly or implicitly, any license or other right to use the information other than for the above-stated training purposes. Copyright © 1999-2018 Bombardier Inc. or its subsidiaries. All rights reserved. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
Acronyms and Abbreviations
ACRONYMS AND ABBREVIATIONS A
AEMF
aft engine mount fitting
AFCS
automatic flight control system
AFCU
alternate flight control unit
A/T
autothrottle
AFDA
adaptive flight display application
ABIT
automatic BIT
AFDT
adaptive flight display table
ABS
absolute
AFDX
avionics full duplex switched Ethernet
ABS
autobrake system
AFM
Airplane Flight Manual
ACARS
aircraft communications addressing and reporting system
AFP
automated fiber placement
ACC
active clearance control
AGB
accessory gearbox
ACES
avionics cooling and extraction system
AGL
above ground level
ACL
access control list
AHC
attitude heading computer
ACM
air cycle machine
AHMS
aircraft health management system
ACM
aircraft condition monitoring
AIS
aircraft information server
ACMF
aircraft condition monitoring function
AIS
audio integrating system
ACMP
AC motor pump
AIM
axle interface module
ACP
audio control panel
AIM
aircraft identification module
ACU
audio conditioning unit
AIM
align-in-motion
ADC
air data computer
AL
autoland
ADEF
aircraft data exchange function
ALC
APU line contactor
ADF
automatic direction finder
ALI
airworthiness limitation item
ADI
attitude direction indicator
AI-Li
aluminum-lithium
ADLS
airborne data link system
ALM
application license manager
ADMF
aircraft data management function
ALT
altitude
ADRF
aircraft data recording function
ALTN FLAP
alternate flap
ADS
air data system
AM
amplitude modulation
ADSP
air data smart probe
AMCU
advanced monitor control unit
ADSP
air data system probe
AMP
Aircraft Maintenance Publication
AES
alternate extension system
ANR
archive noise reduction
AEV
avionics exhaust valve
ANS
aircraft network switch
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
i
Acronyms and Abbreviations
AOA
angle-of-attack
BAHX
buffer air heat exchanger
AOC
air/oil cooler
BAPS
buffer air pressure sensor
AOC
airport operational communications
BAVS
buffer air valve solenoid
AODV
active oil damper valve
BAVSOV
buffer air shutoff valve
AOHX
air/oil heat exchanger
BCCC
base coat clear coat
AP
autopilot
BCS
brake control system
APM
aircraft personality module
BDCU
brake data concentrator unit
APR
automatic power reserve
BFO
beat frequency oscillator
APU
auxiliary power unit
BGM
boarding music
AR
automatic realignment
BIT
built-in test
ARTCC
air route traffic control center
BITE
built-in test equipment
ASA
autoland status annunciator
BL
buttock line
ASC
APU starting contactor
BLC
battery line contactor
ASM
air separation module
BLS
bifurcation latch system
ASRP
aircraft structure repair publication
BMPS
bleed monitoring pressure sensor
AT
autothrottle
BPCU
bus power control unit
ATC
air traffic control
BPMS
bleed pressure monitoring sensor
ATIS
air traffic information services
BSC
battery start contactor
ATP
acceptance test procedure
BTC
bus tie contactor
ATS
air turbine starter
BTS
base transceiver station
ATS
autothrottle system
BTMS
brake temperature monitoring system
ATS
air traffic services
BTS
brake temperature sensor
AV-VENTS
avionics ventilated temperature sensor
BTS
bleed temperature sensor
BVID
barely visible impact damage
B BALODS
bleed air leak and overheat detection system
BAP
buffer air pressure
CADTS
cargo duct temperature sensor
BAS
bleed air system
CAI
cowl anti-ice
BAV
bleed air valve
CAIS
cowl anti-ice system
BACV
buffer air check valve
CAIV
cowl anti-ice valve
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
C
TECHNICAL TRAINING MANUAL
For Training Purposes Only
ii
Acronyms and Abbreviations
CAM
cockpit area microphone
CMUI
configuration management unique identifier
CAN
controller air network
CNS
communications, navigation and surveillance
CAS
calibrated airspeed
COM
communication
CAS
crew alerting system
CPCS
cabin pressure control system
CATS
cargo temperature sensor
CPCV
compensator pressure check valve
CB
circuit breaker
CPD
circuit protection device
CBIT
continuous built-in test
CPDD
circuit protection device detector
CBP
circuit breaker panel
CPDLC
controller-pilot data link communications
CBV
cross-bleed valve
CPN
Collins part number
CC
cabin controller
CPU
central processing unit
CCDL
cross-channel data link
CRC
cyclical redundancy check
CCM
common computing module
CRES
corrosion resistant steel
CCMR
common computing module runtime
CSD
cabin service display
CCP
cursor control panel
CSD
customer service display
CCU
camera control unit
CSMU
crash-survivable memory unit
CCW
counterclockwise
CSOV
cargo shutoff valve
CDC
control and distribution cabinet
CT
crew terminal
CDI
course deviation indicator
CT
current transformer
CDTS
compressor discharge temperature sensor
CTP
control tuning panel
CEM
cover and environmental module
CVR
cockpit voice recorder
CF
configuration file
CW
clockwise
CFIT
controlled flight into terrain
CWB
center wing box
CFRP
carbon fiber reinforced polymer
CIC
compressor intermediate case
D/I
discrete input
CIC
corrosion inhibiting compound
D/O
discrete output
CLAWS
control laws
DBM
database manager
CM
configuration manager
DCM
data concentrator module
CMS
cabin management system
DCMR
data concentration module runtime
CMU
communication management unit
DCS
data concentration system
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
D
TECHNICAL TRAINING MANUAL
For Training Purposes Only
iii
Acronyms and Abbreviations
DCU
directional control unit
EFAN
extraction fan
DCV
directional control valve
EDM
emergency descent mode
DFSOV
dual-flow shutoff valve
EDP
engine-driven hydraulic pump
DLCA
data link communications application
EDP
engine-driven pump
DMA
display manager application
EDU
electronic display unit
DMC
data concentrator unit module cabinet
EEC
electronic engine control
DME
distance measuring equipment
EEGS
emergency electrical power generation
DMM
data memory module
EEPROM
electrical erasable programmable read only memory
DP
differential protection
EESS
emergency escape slide system
DPCT
differential protection current transformer
EFB
electronic flight bag
DPI
differential pressure indicator
EEGS
emergency electrical power generation
DPLY
deploy
EFCS
electronic flight control system
DRA
diagnostic and reporting application
eFIM
electronic fault isolation manual
DSK
double-stack knob
EFIS
electronic flight instrument system
DSM
digital switching module
EGT
exhaust gas temperature
DSU
data storage unit
EHSV
electrohydraulic servovalve
DSPU
diode shunt protection unit
EIC
engine inlet cowl
DTC
DC tie contactor
EICAS
engine indication and crew alerting system
DTE
damage tolerance evaluation
ELC
external power line contactor
DTI
damage tolerance inspection
ELT
emergency locator transmitter
DTS
duct temperature sensor
EMA
electric motor actuator
DU
display unit
EMU
expansion module unit
E
EMAC
electric motor actuator controller
EBC
essential bus contactor
EMCU
electric motor control unit
ECL
electronic checklist
EMER
emergency
ECS
environmental control system
EMPC
emergency power control
ECU
electronic control unit
EOAM
emergency opening assist means
ECU
external compensation unit
EOF
end-of-flight
EDCM
electronic door control module
EPC
electrical power center
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
iv
Acronyms and Abbreviations
EPC
electronic power center
FCV
flow control valve
EPGD
electrical power generation and distribution
FD
flight director
EPGDS
electrical power generation and distribution system
FDDSS
flight deck door surveillance system
EPGS
electrical power generation system
FDE
flight deck effect
EPP
engine programming plug
FDG
fan drive gearbox
EPSU
emergency power supply unit
FDGS
fan drive gear system
EPTRU
external power transformer rectifier unit
FDR
flight data recorder
ERAV
emergency ram air valve
FDRAS
flight deck remote access system
ESD
electrostatic discharge
FDV
flow divider valve
ETC
essential tie contactor
FEGV
fan exit guide vane
ETOPS
extended-range twin-engine operational performance standards
FEMB
forward engine mount bulkhead
FF
fuel flow
F
FFDP
fuel flow differential pressure
FA
flight attendant
FFSV
free fall selector valve
FAA
Federal Aviation Authority
FG
flight guidance
FADEC
full authority digital engine control
FGS
flight guidance system
FANS
future air navigation system
FIC
fan intermediate case
FAV
fan air valve
FIDEX
fire detection and extinguishing
FBC
front bearing compartment
FIM
fault isolation manual
FBW
fly-by-wire
FLC
flight level change
FBWPC
fly-by-wire power converter
FLS
fast load-shed
FC
full close
FLTA
forward-looking terrain avoidance
FCP
flight control panel
FMA
flight mode annunciator
FCS
flight control system
FMS
flight management system
FCBS
fatigue critical baseline structure
FMSA
flight management system application
FCEE
flight crew emergency exit
FMV
fuel metering valve
FCSB
fan cowling support beam
FO
full open
FCU
flush control unit
FOD
foreign object debris
FCU
fuel control unit
FOHX
fuel/oil heat exchanger
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
v
Acronyms and Abbreviations
FOHXBV
fuel/oil heat exchanger bypass valve
GSA
ground spoiler actuator
FPS
feedback position sensor
GSCM
ground spoiler control module
FPV
flight path vector
GSE
ground support equipment
FQC
fuel quality computer
GUI
graphical user interface
FS
fixed structure
FS
flight station
HAAO
high altitude airport operation
FS
fuselage station
HDG
heading
FSA
file server application
HF
high frequency
FSV
flow sensor venturi
HID
high-intensity discharge
FSB
fasten seat belt
HLEIF
high load event indication function
FSCL
flight spoiler control lever
HLSL
high lift selector lever
FTIS
fuel tank inerting system
HMU
health management unit
FW
failure warning
HOR
hold open rod
FWSOV
firewall shutoff valve
HP
high-pressure
HPC
high-pressure compressor
H
G GA
go-around
HPD
hydraulic pump depressurization
GCF
ground cooling fan
HPGC
high-pressure ground connection
GCR
generator control relay
HPSOV
high-pressure shutoff valve
GCS
global connectivity suite
HPT
high-pressure turbine
GCU
generator control unit
HPV
high-pressure valve
GFP
graphical flight planning
HRD
high-rate discharge
GFRP
glass fiber reinforced polymer
HRTDb
high-resolution terrain database
GHTS
galley heater temperature sensor
HS
handset
GLC
generator line contactor
HS
high solid
GMT
Greenwich mean time
HSI
horizontal situation indicator
GNSS
global navigation satellite system
HSTA
horizontal stabilizer trim actuator
GPWS
ground proximity warning system
HSTS
horizontal stabilizer trim system
GS
glideslope
HUD
head-up display
GS
ground spoiler
HUDS
head-up display system
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
vi
Acronyms and Abbreviations
I
IOM
input/output module
I/O
input/output
IP
information provider
IAMS
integrated air management system
IPC
integrated processing cabinet
IAS
indicated airspeed
IPCKV
intermediate pressure check valve
IASC
integrated air system controller
IPS
inches per second
IBIT
initiated built-in test
IPS
integrated processing system
IBR
integral bladed rotor
IRCV
inlet return check valve
ICAO
International Civil Aviation Organization
IRS
inertial reference system
ICCP
integrated cockpit control panel
IRU
inertial reference unit
ICDU
integrated control display unit
ISI
integrated standby instrument
ICU
inerting control unit
ISM
input signal management
ICU
isolation control unit
ISPS
in-seat power supply
ICV
isolation control valve
ISPSS
in-seat power supply system
IFE
in-flight entertainment system
ITT
interturbine temperature
IFEC
in-flight entertainment and connectivity system
IFIS
integrated flight information system
IFPC
integrated fuel pump and control
IFS
information landing system
L/S
lube/scavenge
IFS
inner fixed structure
LAN
local area network
IGN
ignition
LBIT
landing built-in test
IGV
inlet guide vane
LCD
life cycle data
IGVA
inlet guide vane actuator
LCT
line current transformer
IIM
inceptor interface module
LED
light-emitting diode
IIV
inlet isolation valve
LLU
LED lighting unit
ILS
instrument landing system
LGCL
landing gear control lever
IMA
integrated modular avionics
LGCV
landing gear control valve
IMS
information management system
LGIS
landing gear indicating system
INT
intermittent
LGSCU
landing gear and steering control unit
IOC
input/output concentrator
LGSV
landing gear selector valve
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
J JOSV
journal oil shuttle valve L
TECHNICAL TRAINING MANUAL
For Training Purposes Only
vii
Acronyms and Abbreviations
LOC
localizer
MIXTS
mix manifold temperature sensor
LOP
low oil pressure
MKP
multifunction keyboard panel
LOPA
layout of passenger area
MLG
main landing gear
LP
low-pressure
MLW
maximum landing weight
LPC
low-pressure compressor
MMEL
Master Minimum Equipment List
LPSOV
low-pressure shutoff valve
MOF
main oil filter
LPT
low-pressure turbine
MOT
main oil temperature
LRD
low-rate discharge
MPP
maintenance planning publication
LRM
line replaceable module
MPSOV
minimum pressure-shutoff valve
LRU
line replaceable unit
MRW
maximum ramp weight
LSK
line select key
MSV
mode select valve
LSOP
lubrication and scavenge oil pump
MTD
master time and date
LV
lower sideband voice
MTO
maximum rated takeoff
LVDS
low-voltage differential signaling
MTOW
maximum takeoff weight
LVDT
linear variable differential transformer
MWW
main wheel well
MZFW
maximum zero fuel weight
M MAX
maximum
MB
marker beacon
NA
not activated
MCDL
motor control data link
NACA
National Advisory Committee for Aeronautics
MCE
motor control electronic
ND
nosedown
MCR
minimum control requirement
NCD
no computed data
MCV
mode control valve
NCG
network communication gap
MDU
manual drive unit
NCU
network control unit
MEL
minimum equipment list
NDB
non-directional beacon
MES
main engine start
NDO
network data object
MFK
multifunction keyboard panel
NEA
nitrogen-enriched air
MFP
multifunction probe
NEADS
nitrogen-enriched air distribution system
MFS
multifunction spoiler
NLG
nose landing gear
MFW
multifunction window
NO PED
no personal electronic device
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
N
TECHNICAL TRAINING MANUAL
For Training Purposes Only
viii
Acronyms and Abbreviations
NPRV
negative pressure-relief valve
NU
noseup
P&W
Pratt and Whitney
NWS
nosewheel steering
P2
inlet pressure
NVM
non-volatile memory
PA
passenger address
O
PAX
passenger
OAT
outside air temperature
PBA
pushbutton annunciator
OBB
outboard brake
PBE
protective breathing equipment
OBIGGS
onboard inlet gas generation system
PBIT
power-up built-in test
OC
overcurrent
PCE
precooler exhaust
OCM
oil control module
PCE
precooler exit
OCM
option control module
PCU
power control unit
ODI
overboard discharge indicator
PDF
portable document format
ODL
onboard data loader
PDOS
power door operating system
ODM
oil debris monitor
PDPS
pack pressure differential sensor
OEA
oxygen-enriched air
PDL
permitted damage limits
OEM
original equipment manufacturer
PDS
power distribution system
OF
overfrequency
PDTS
pack discharge temperature sensor
OFV
outflow valve
PDU
power drive unit
OMS
onboard maintenance system
PED
personal electronic device
OMS IMA
OMS interactive maintenance application
PEM
power environment module
OMSA
onboard maintenance system application
PEV
pressure equalization valve
OMST
onboard maintenance system table
PFCC
primary flight control computer
OPAS
outboard position asymmetry sensor
PFD
primary flight display
OPU
overvoltage protection unit
PFS
post flight summary
OSP
opposite-side pressure
PHMU
prognostics and health management unit
OSS
overspeed/shutdown solenoid
PIC
peripheral interface controller
OT
other traffic
PIC
processor-in-command
OV
overvoltage
PIFS
pack inlet flow sensor
OWEE
overwing emergency exit
PIM
panel interface module
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
P
TECHNICAL TRAINING MANUAL
For Training Purposes Only
ix
Acronyms and Abbreviations
PIPS
pack inlet pressure sensor
PVT
position, velocity, time
PLD
programmable logic device
PWM
pulse width modulation
PLD
proportional lift dump
PMA
permanent magnet alternator
QAD
quick attach/detach
PMA
program manager application
QEC
quick engine change
PMAG
permanent magnet alternator generator
PMG
permanent magnet generator
RA
radio altimeter
POB
power off brake
RA
resolution advisory
POB
pressure off brake
RAM
receiver autonomous integrity monitoring
POR
point of regulation
RAD
radio altitude
PPM
power producing module
RARV
ram air regulating valve
PPT
pedal position transducer
RAT
ram air turbine
PRAM
prerecorded announcement and message
RDC
remote data concentrator
PRSOV
pressure-regulating shutoff valve
RDCP
refuel/defuel control panel
PRV
pressure-regulating valve
REL
relative altitude
PRV
pressure-relief valve
REO
repair engineering order
PS
passenger service
RET
retracted
PS
pressure sensor
REU
remote electronic unit
PSA
print server application
RF
radio frequency
PSE
principal structural element
RFAN
recirculation fan
PSU
passenger service unit
RGA
rotary geared actuator
PSUC
passenger service unit controller
RGC
RAT generator control
PT
proximate traffic
RIPS
recorder independent power supply
PT
pressure transducer
RIU
radio interface unit
Pt
total pressure
RLC
RAT line contactor
PTS
pack temperature sensor
RMA
remote maintenance access
PTT
push-to-talk
RMS
radio management system
PTU
power transfer unit
ROLS
remote oil lever sensor
PTY
priority
ROV
redundant overvoltage
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Q
R
TECHNICAL TRAINING MANUAL
For Training Purposes Only
x
Acronyms and Abbreviations
RPA
rudder pedal assembly
SFV
safety valve
RPM
revolutions per minute
SLS
slow load-shed
RSA
report server application
SMS
surface management system
RSP
reversion switch panel
SOV
solenoid operated valve
RTA
receiver-transmitter antenna
SOV
shutoff valve
RTD
resistance temperature device
SPCV
supply pressure check valve
RTD
resistive thermal device
SPDS
secondary power distribution system
RTL
ready-to-load
SPDT
single pole double throw
RTO
rejected takeoff
SPKR
speaker
RTS
return to service
SPM
seat power module
RTSA
radio tuning system application
SSC
sidestick controller
RVDT
rotary variable differential transformer
SSD OML
solid-state onboard media loader
SSI
structural significant item
S SAL
specific airworthiness limitation
SSEC
static source error connection
SAT
static air temperature
SSPC
solid-state power controller
SATCOM
satellite communication
SSPC-CB
solid-state power controller circuit breaker
SAV
starter air valve
SSRPC
solid-state remote power controller
SB
service bulletin
SUA
special use airspace
SBAS
satellite-based augmentation system
SVA
stator vane actuator
SBIT
start-up BIT
SVS
synthetic vision system
SCV
surge control valve
SCV
steering control valve
T/M
torque motor
SEB
seat electronics box
T/R
thrust reverser
SELCAL
selective calling
T2
inlet temperature
SFCC
slat/flap control computer
TA
traffic advisory
SFCL
slat/flap control lever
TACKV
trim air check valve
SFCP
slat/flap control panel
TAPRV
trim air pressure-regulating valve
SFECU
slat/flap electronic control unit
TASOV
trim air shutoff valve
SFIS
standby flight instrument system
TAT
total air temperature
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
T
TECHNICAL TRAINING MANUAL
For Training Purposes Only
xi
Acronyms and Abbreviations
TAV
trim air valve
TQA
throttle quadrant assembly
TAWS
terrain awareness and warning system
TRAS
thrust reverser actuation system
TAWSDb
terrain awareness and warning system database
TRU
transformer rectifier unit
TCA
turbine cooling air
TSC
TRU start contactor
TCAS
traffic alert and collision avoidance system
TSFC
thrust specific fuel consumption
TCB
thermal circuit breaker
TSM
trip status monitor
TCDS
type certificate data sheet
TSO
technical standard order
TCF
terrain control valve
TSS
traffic surveillance system
TDR
transponder
TTG
time-to-go
TCV
temperature control valve
TTP
time-triggered protocol
TEC
turbine exhaust case
TWIP
terminal weather information for pilot
TED
trailing edge down
TEU
trailing edge up
UART
universal asynchronous receiver transmitter
TFTP
trivial file transfer protocol
UBMF
usage-based monitoring function
TIC
turbine inlet case
UF
underfrequency
TIC
turbine intermediate case
ULB
underwater locator beacon
TIV
temperature inlet valve
UPLS
ultrasonic point level sensors
TIV
temperature isolation valve
USB
universal serial bus
TLA
throttle lever angle
UTC
universal time coordinated
TLC
TRU line contactor
UV
upper sideband voice
TLD
time limited dispatch
UV
ultraviolet
TOGA
takeoff/go-around
UV
undervoltage
TPIS
tire pressure indicating system
TPM
TAWS processing module
VAC
volts alternating current
TPM
tire pressure module
VDC
voltage direct current
TPMA
terrain processing module application
VDL
VHF data link
TPMU
tire pressure monitoring unit
VDLM
VHF data link mode
TPS
tire pressure sensor
VENTS
ventilated temperature sensor
TPSA
terrain processing system application
VFG
variable frequency generator
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
U
V
TECHNICAL TRAINING MANUAL
For Training Purposes Only
xii
Acronyms and Abbreviations
VFGOOHX
variable frequency generator oil/oil heat exchanger
WST
wheel speed transducer
VGMD
vacuum generator motor drive unit
WTBF
wing-to-body fairing
VHF-NAV
VHF navigation
WWHS
windshield and side window heating system
VID
visible impact damage
WWS
waste water system
VL
virtual link
WWSC
water and waste system controller
VLAN
virtual local area network
WXR
weather radar
VNAV
vertical navigation
VOC
volatile organic compounds
ZB
zone box
VOR-VHF
VHF omnidirectional radio
P
differential pressure
VORV
variable oil reduction valve
VPA
video passenger announcement
VSD
vertical situation display
VSPD
V-speed
VSWR
voltage standing-wave radio
VTU
video transmission unit
Z
W WAI
wing anti-ice
WAP
wireless access point
WAIS
wing anti-ice system
WAITS
wing anti-ice temperature sensor
WAIV
wing anti-ice valve
WBV
windmill bypass valve
WIPC
windshield ice protection controller
WL
waterline
WOFFW
weight-off-wheels
WOW
weight-on-wheels
WPS
words-per-second
WS
wing situation
WSA
web server application
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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Acronyms and Abbreviations
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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xiv
Module 5: Power Plant - List of Changes
MODULE 5: POWER PLANT - LIST OF CHANGES The following table details the changes applied to this revision: ATA NAME AND NUMBER
CMUI NUMBER
CHANGES APPLIED
Front Page
CS130-21.07-06-00-121418
Version number and CMUI number
Acronyms and Abbreviations
CS130-21.07-06-00-121418
Deleted BCT (bus tie contactor current transformer) and added MKP (multifunction keyboard panel) (pges ii and viii)
List of Changes
CS130-21.07-06-00-121418
New CMUI
21 Environmental Control
CS130-21.07-06-00-121418
No change
28 Fuel
CS130-21.07-06-00-121418
Added new Fuel Safety section (pges 28-2 to 28-13) Updated collector tank area (pge 28-21) Updated graphic (pges 28-31, 28-33, 28-35) Added maximum suction pressure (pge 28-36) Updated graphic (pges 28-37, 28-39, 28-71, 28-87, 28-101, 28-103, 28-105) Added caution (pge 28-106) Updated graphic (pges 28-107, 28-109, 28-111)
30 Ice and Rain Protection
CS130-21.07-06-00-121418
Moved System Test before CAS Messages (pges 30-30 to 30-35)
36 Pneumatic
CS130-21.07-06-00-121418
No change
47 Liquid Nitrogen (Fuel Tank Inerting System)
CS130-21.07-06-00-121418
No change
49 Airborne Auxiliary Power
CS130-21.07-06-00-121418
Updated graphic (pges 49-9, 49-13) Added note and caution (pge 49-10) Updated text (pge 49-14) Updated graphic (pges 49-15, 49-23, 49-97)
71 Power Plant
CS130-21.07-06-00-121418
No change
72 Engine
CS130-21.07-06-00-121418
Updated graphic (pge 72-3)
73 Engine Fuel and Control
CS130-21.07-06-00-121418
Changed throttle lever angle to thrust lever angle (pge 73-43)
74-80 Ignition and Starting
CS130-21.07-06-00-121418
No change
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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LOC i
Module 5: Power Plant - List of Changes
ATA NAME AND NUMBER
CMUI NUMBER
CHANGES APPLIED
75 Engine Air Systems
CS130-21.07-06-00-121418
No change
76 Engine Controls
CS130-21.07-06-00-121418
No change
77 Engine Indicating
CS130-21.07-06-00-121418
Updated graphic; added indicator for fan drive shaft dimple location, and view forward looking aft (pge 77-43) Added new Fan Trim Balance Using the PW 1500G Trim Balance Helper section with graphics (pges 77-58 to 77-69)
78 Exhaust
CS130-21.07-06-00-121418
Updated graphic (pge 78-41) Updated text (pge 78-42) Changed ICV control solenoid to ICU isolation valve control solenoid (pge 78-44) Changed valve to unit in figure title (pge 78-45) Changed DCV energized to DCU energized in heading and figure title (pges 78-48, 78-49) Changed DCV control solenoid to DCU control solenoid (pge 78-52) Changed ICV control solenoid to ICU control solenoid (pge 78-54) Changed ICV and DCV enable solenoids to ICU and DCU control solenoids; ICV solenoid to ICU solenoid; DCV solenoid to DCU solenoid; added information for maximum reverse thrust (pge 78-56)
79 Oil
CS130-21.07-06-00-121418
Added text for oil level and updated caution (pge 79-4) Updated graphic (pge 79-23) Updated tables in graphic (pge 79-37) Changed 47°C (117°F) to 49°C (120°F) (pge 79-38) Updated EICAS on graphic (pge 79-39) Updated graphic (pge 79-41)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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LOC ii
ATA 21 - Environmental Control
BD-500-1A10 BD-500-1A11
21 - Environmental Control
Table of Contents 21-90 Integrated Air System Controller ..................................21-2
21-51 Cooling....................................................................... 21-34
General Description ..........................................................21-2 Component Location .......................................................21-4 Detailed Description .........................................................21-6 Monitoring and Tests .......................................................21-8 INFO Messages ...........................................................21-9 Practical Aspects.............................................................21-10 Onboard Maintenance System Test ..........................21-10
General Description ...................................................... 21-34 Component Location .................................................... 21-36 Water Extractor ......................................................... 21-38 Water Sprayer ........................................................... 21-38 Pack Discharge Pressure Sensor ............................. 21-40 Component Information ................................................ 21-42 Temperature Control Valve ....................................... 21-42 Ram Air Regulating Valve ........................................ 21-44 Controls and Indications ................................................ 21-46 Controls..................................................................... 21-46 Indications ................................................................. 21-48 Detailed Description ..................................................... 21-50 Pack Temperature Regulation .................................. 21-50 Pack Icing Protection ............................................... 21-50 Pack Overheat Protection ........................................ 21-50 Pack Discharge Overheat Protection ....................... 21-50 Monitoring and Tests .................................................... 21-52 CAS Messages ......................................................... 21-53
21-53 Flow Control System ..................................................21-12 General Description .......................................................21-12 Component Location .....................................................21-14 Flow Sensor Venturi...................................................21-14 Flow Control Valve.....................................................21-14 Ozone Converter/Ozone and VOC Converter (Optional)..........................................21-14 Pack Inlet Pressure Sensors......................................21-16 Pack Inlet Flow Sensors ............................................21-16 Controls and Indications ................................................21-18 Controls......................................................................21-18 Indications..................................................................21-20 Detailed Description .......................................................21-22 Flow Control Normal Operation .................................21-22 Flow Control...............................................................21-24 Flow Control Modes ...................................................21-26 Flow Control Shutoff ..................................................21-28 Flow Control Valve Failure.........................................21-28 Monitoring and Tests ......................................................21-30 CAS Messages ..........................................................21-31 Practical Aspects.............................................................21-32
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
21-52 Ram Air System......................................................... 21-56 General Description ...................................................... 21-56 Component Location .................................................... 21-58 Emergency Ram Air Valve ........................................ 21-58 Controls and Indications ................................................ 21-60 Detailed Description ..................................................... 21-62 Monitoring and Tests .................................................... 21-64 CAS Messages ......................................................... 21-65 21-60 Trim Air System ......................................................... 21-66 General Description ...................................................... 21-66
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21 - Environmental Control
Component Location .....................................................21-68 Trim Air Valves...........................................................21-70 Controls and Indications ................................................21-72 Detailed Description ......................................................21-74 Trim Air Control..........................................................21-74 Monitoring and Tests ......................................................21-76 CAS Messages ..........................................................21-77 21-60 Temperature Control System .....................................21-78 General Description .......................................................21-78 Component Location ......................................................21-80 Mix Manifold Temperature Sensors ..........................21-80 Duct Temperature Sensors .......................................21-82 Ventilated Temperature Sensors ..............................21-84 Cargo Duct Temperature Sensor...............................21-86 Cargo Temperature Sensor ......................................21-86 Component Information ..................................................21-88 Ventilated Temperature Sensor .................................21-88 Controls and Indications .................................................21-90 AIR Panel...................................................................21-90 AIR Synoptic Page.....................................................21-92 Cabin Management System Temperature Controls ................................................21-94 Detailed Description ......................................................21-96 Automatic Temperature Control.................................21-96 Manual Temperature Control .....................................21-96 FWD Cargo Temperature Control..............................21-98 Monitoring and Tests ....................................................21-100 CAS Messages ........................................................21-101 21-22 Recirculation.............................................................21-102 General Description .....................................................21-102 Component Location ...................................................21-104 Component Information ...............................................21-106 Recirculation Filters .................................................21-106 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Controls and Indications ............................................. 21-108 Detailed Description ................................................... 21-110 Monitoring and Test .................................................... 21-112 CAS Messages ....................................................... 21-113 21-22 Distribution System.................................................. 21-114 General Description .................................................... 21-114 Zone Distribution ..................................................... 21-116 Overhead Bin Distribution ....................................... 21-120 Flight Deck Distribution ........................................... 21-122 Component Location .................................................. 21-124 Component Information .............................................. 21-126 Low-Pressure Ground Connection.......................... 21-126 Mix Manifold ............................................................ 21-128 21-23 Ventilation System ................................................... 21-130 General Description .................................................... 21-130 Component Location .................................................. 21-132 Controls and Indications .............................................. 21-134 AIR Panel ................................................................ 21-134 Cargo Panel ............................................................ 21-134 Indications ............................................................... 21-136 Detailed Description .................................................... 21-138 Monitoring and Tests ................................................... 21-140 CAS Messages ....................................................... 21-141 Practical Aspects ......................................................... 21-142 21-40 Heating .................................................................... 21-144 Flight Crew Floor Heating ................................................. 21-144 General Description ..................................................... 21-144 Component Location .................................................. 21-144 Heated Mat.............................................................. 21-144 Detailed Description ................................................... 21-146
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21 - Environmental Control
Galley Heating ...................................................................21-148 General Description ......................................................21-148 Component Location ...................................................21-150 Galley Fan................................................................21-150 Galley Heater ...........................................................21-150 Galley Heater Temperature Sensor .........................21-150 Controls and Indications ..............................................21-152 Galley Heat Controls................................................21-152 Galley Heat Indications ...........................................21-154 Detailed Description ....................................................21-156 21-26 Avionics Cooling and Extraction System ..................21-158 General Description .....................................................21-158 Air Inlet Cooling System ....................................................21-160 General Description .....................................................21-160 Component Location ...................................................21-162 Forward Avionics Filter ............................................21-162 Forward Bay Fan .....................................................21-162 Forward Skin Heat Exchanger .................................21-162 Ground Valve ...........................................................21-162 Mid Bay Fan.............................................................21-162 Avionics Coalescent Filter........................................21-162 Mid Skin Heat Exchanger ........................................21-162 Mid Avionics Filter....................................................21-162 Ducting.....................................................................21-162 Component Information ................................................21-164 Forward Avionics Filter ............................................21-164 Forward Skin Heat Exchanger .................................21-164 Ground Valve ...........................................................21-166 Avionics Coalescent Filter........................................21-166 Mid Skin Heat Exchanger ........................................21-168 Mid Avionics Filter....................................................21-168 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Controls and Indications ............................................. 21-170 Detailed Description ................................................... 21-172 Air Extraction System ........................................................ 21-174 General Description .................................................... 21-174 Component Location ................................................... 21-176 Extraction Fans ....................................................... 21-176 Ducting .................................................................... 21-176 Detailed Description ................................................... 21-178 Backup Exhaust and Extraction System ........................... 21-180 General Description .................................................... 21-180 Component Location .................................................. 21-182 Avionics Exhaust Valves ......................................... 21-182 Backup Fan ............................................................. 21-182 Ducting .................................................................... 21-182 Avionics Bay Ventilated Temperature Sensors ....... 21-184 Detailed Description .................................................... 21-186 Component Information .............................................. 21-188 Avionics Exhaust Valves ......................................... 21-188 Backup Fan ............................................................. 21-190 Controls and Indications ............................................. 21-192 Monitoring and Tests .................................................. 21-194 CAS Messages ....................................................... 21-195 Wing-to-Body Fairing Ventilation System ......................... 21-198 General Description ..................................................... 21-198 High-Pressure Pneumatic Duct Area Cooling ......... 21-200 21-30 Cabin Pressure Control System .............................. 21-202 General Description .................................................... 21-202 Automatic Mode ...................................................... 21-202 Manual Mode .......................................................... 21-202 Ditching ................................................................... 21-202
TECHNICAL TRAINING MANUAL
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CS130_M5_21_EnviroContTOC.fm
21 - Environmental Control
Emergency Depressurization...................................21-202 Safety Functions ......................................................21-202 Typical Flight Profile.................................................21-204 Component Location ...................................................21-206 Outflow Valve...........................................................21-206 Water Activated Check Valve ..................................21-206 Muffler ......................................................................21-206 Safety Valve.............................................................21-206 Negative Pressure-Relief Valve ...............................21-206 Pressure Equalization Valves ..................................21-206 Component Information ...............................................21-208 Outflow Valve...........................................................21-208 Water Activated Check Valve .................................21-210 Muffler ......................................................................21-210 Safety Valve.............................................................21-212 Negative Pressure-Relief Valve ...............................21-212 Pressure Equalization Valves ..................................21-214 Detailed Component Information .................................21-216 Integrated Air System Controller..............................21-216 Controls and Indications ..............................................21-218 Pressurization Panel ................................................21-218 FMS Page ................................................................21-220 Indications................................................................21-222 Detailed Description ....................................................21-224 Auto/Manual Selection .............................................21-224 Interface ...................................................................21-226 Cabin Altitude Limitation Annunciation ....................21-228 Emergency Depressurization...................................21-230 Ditching ....................................................................21-232 Monitoring and Tests ...................................................21-234 CAS Messages ........................................................21-235
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CS130_M5_21_EnviroContTOC.fm
21 - Environmental Control
List of Figures Figure 1: Integrated Air System Controller ...........................21-3 Figure 2: Integrated Air System Controller Location.............21-5 Figure 3: Integrated Air System Controller Detailed Description ..............................................21-7 Figure 4: Environmental Control System OMS Page .........21-11 Figure 5: Flow Control System ...........................................21-13 Figure 6: Flow Sensor Venturi, Flow Control Valve, and Ozone Converter Location ...........................21-15 Figure 7: PIP Sensor and PIF Sensor Locations................21-17 Figure 8: Flow Control Controls..........................................21-19 Figure 9: Flow Control Indications ......................................21-21 Figure 10: Flow Control Normal Operation...........................21-23 Figure 11: Flow Control Valve ..............................................21-25 Figure 12: Flow Control Priority ............................................21-27 Figure 13: Flow Control Shutoff............................................21-29 Figure 14: Flow Control Valve Lockout.................................21-33 Figure 15: Air Conditioning System ......................................21-35 Figure 16: Air Conditioning Components Location ...............21-37 Figure 17: Water Extractor and Water Sprayer ....................21-39 Figure 18: Pack Discharge Pressure Sensor Location........21-41 Figure 19: Temperature Control Valve .................................21-43 Figure 20: Ram Air Regulating Valve ...................................21-45 Figure 21: Air Conditioning Pack Controls............................21-47 Figure 22: Air Conditioning Pack Indications........................21-49 Figure 23: Air Conditioning System Detailed Description ............................................21-51 Figure 24: Ram Air System ..................................................21-57 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 25: Emergency Ram Air Valve.................................. 21-59 Figure 26: Ram Air System Controls and Indications .......... 21-61 Figure 27: Ram Air System Schematic ................................ 21-63 Figure 28: Trim Air System .................................................. 21-67 Figure 29: Trim Air Shutoff Valves, Trim Air Check Valves, and Trim Air Pressure-Regulating Valve Location..... 21-69 Figure 30: Trim Air Valve Locations..................................... 21-71 Figure 31: Trim Air System Controls and Indications........... 21-73 Figure 32: Trim Air Control................................................... 21-75 Figure 33: Temperature Control System.............................. 21-79 Figure 34: Mix Manifold Temperature Sensors.................... 21-81 Figure 35: Duct Temperature Sensors................................. 21-83 Figure 36: Flight Deck and Cabin Ventilated Temperature Sensors ......................................... 21-85 Figure 37: Cargo Duct Temperature Sensor and Cargo Temperature Sensor ......................... 21-87 Figure 38: Ventilated Temperature Sensor.......................... 21-89 Figure 39: Temperature System Controls............................ 21-91 Figure 40: Temperature System Indications ........................ 21-93 Figure 41: Cabin Management System Temperature Controls......................................... 21-95 Figure 42: Flight Deck, Forward Cabin, and Aft Cabin Temperature Control System.............. 21-97 Figure 43: Cargo Temperature Control................................ 21-99 Figure 44: Recirculation System ........................................ 21-103 Figure 45: Recirculation System Component Location...... 21-105
TECHNICAL TRAINING MANUAL
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CS130_M5_21_EnviroContLOF.fm
21 - Environmental Control
Figure 46: Recirculation Filter.............................................21-107 Figure 47: Recirculation System Controls and Indications....................................21-109 Figure 48: Recirculation Fan Detailed Description .............21-111 Figure 49: Distribution System General Description...........21-115 Figure 50: Zone Distribution CS100 ...................................21-117 Figure 51: Zone Distribution CS300 ...................................21-119 Figure 52: Overhead Bin Distribution .................................21-121 Figure 53: Flight Deck Distribution .....................................21-123 Figure 54: Distribution System Component Location .........21-125 Figure 55: Low-Pressure Ground Connection ....................21-127 Figure 56: Mix Manifold ......................................................21-129 Figure 57: Ventilation System Component Location ..........21-131 Figure 58: Cargo Shutoff Valves Component Location ......21-133 Figure 59: Ventilation System Controls ..............................21-135 Figure 60: Ventilation System Controls and Indications .....21-137 Figure 61: Ventilation System Detailed Description ...........21-139 Figure 62: Cargo Shutoff Valve Practical Aspects..............21-143 Figure 63: Flight Crew Floor Heating General Description...........................................21-145 Figure 64: Flight Crew Floor Heating Detailed Description ..........................................21-147 Figure 65: Galley Heating...................................................21-149 Figure 66: Supplemental Galley Heat Component Location .........................................21-151 Figure 67: Galley Heater Controls ......................................21-153 Figure 68: Galley Heater Indications ..................................21-155 Figure 69: Supplemental Galley Heat Detailed Description ..........................................21-157 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 70: Avionics Cooling and Extraction System General Description .......................................... 21-159 Figure 71: Air Inlet Cooling Schematic............................... 21-161 Figure 72: Air Inlet Cooling Component Location .............. 21-163 Figure 73: Forward Avionics Filter and Forward Skin Heat Exchanger.......................... 21-165 Figure 74: Ground Valve and Avionics Coalescent Filter............................................... 21-167 Figure 75: Mid Skin Heat Exchanger and Mid Avionics Filter............................................. 21-169 Figure 76: Avionics Cooling and Extraction System Controls .............................. 21-171 Figure 77: Air Inlet Cooling System ................................... 21-173 Figure 78: Air Extraction System ....................................... 21-175 Figure 79: Main Air Extraction System Component Location......................................... 21-177 Figure 80: Main Air Extraction System............................... 21-179 Figure 81: Backup Exhaust and Air Extraction System General Description .......................................... 21-181 Figure 82: Backup Exhaust and Air Extraction System .... 21-183 Figure 83: Avionics Bay Ventilated Temperature Sensor Location................................................ 21-185 Figure 84: Backup Exhaust and Air Extraction System Schematic ......................................................... 21-187 Figure 85: Avionics Exhaust Valve .................................... 21-189 Figure 86: Backup Fan....................................................... 21-191 Figure 87: Avionics Cooling and Extraction System Controls and Indications ................................... 21-193 Figure 88: Wing-to-Body Fairing Ventilation System ......... 21-199
TECHNICAL TRAINING MANUAL
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CS130_M5_21_EnviroContLOF.fm
21 - Environmental Control
Figure 89: High-Pressure Pneumatic Duct Area Cooling ..............................................................21-201 Figure 90: Cabin Pressure Control System ........................21-203 Figure 91: Typical Flight Profile ..........................................21-205 Figure 92: Pressurization Components Location................21-207 Figure 93: Outflow Valve ....................................................21-209 Figure 94: Outflow Valve, Water Activated Check Valve, and Muffler ........................................................21-211 Figure 95: Safety Valves and Negative Pressure-Relief Valves......................................21-213 Figure 96: Pressure Equalization Valves............................21-215 Figure 97: Integrated Air System Controller .......................21-217 Figure 98: Pressurization Panel .........................................21-219 Figure 99: FMS, Control Tuning Panel, and AVIONIC CTP Controls ..............................21-221 Figure 100:EICAS and AIR Synoptic Page Indications ......21-223 Figure 101:Cabin Pressure Control System Selection .......21-225 Figure 102:Cabin Pressure Control System Detailed Description ..........................................21-227 Figure 103:Cabin Altitude Limitations Annunciation ...........21-229 Figure 104:Emergency Depressurization ...........................21-231 Figure 105:Ditching.............................................................21-233
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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ENVIRONMENTAL CONTROL - CHAPTER BREAKDOWN ENVIRONMENTAL CONTROL CHAPTER BREAKDOWN
Air Conditioning
1 Avionics Cooling and Extraction System
2 Cabin Pressure Control System
3
21 - Environmental Control 21-90 Integrated Air System Controller
21-90 INTEGRATED AIR SYSTEM CONTROLLER GENERAL DESCRIPTION Control and monitoring of the environmental control system (ECS) is provided by two integrated air system controllers (IASCs). The environmental control system includes: •
Air conditioning system
•
Avionics cooling and extraction system
•
Cabin pressure control system
The IASCs provide system information to the engine indication and crew alerting system (EICAS), the AIR synoptic page, and the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_IASC_General.fm
21 - Environmental Control 21-90 Integrated Air System Controller
Pneumatic System
BAS
BALODS
EICAS
IASC 1
SYN OMS
WAIS
Environmental Control System IASC 2 ACS
CPCS
ACES ACS CPCS IASC
Bleed Air Cold Air Avionics Cooling and Extraction System Air Conditioning System Cabin Pressure Control System Integrated Air System Controller
ACES
CS1_CS3_2190_004
LEGEND
Figure 1: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_IASC_General.fm
21 - Environmental Control 21-90 Integrated Air System Controller
COMPONENT LOCATION The IASCs are located in the mid equipment bay on the lower forward shelf.
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21 - Environmental Control 21-90 Integrated Air System Controller
IASC 2
IASC 1
INTEGRATED AIR SYSTEM CONTROLLER
CS1_CS3_2190_002
MID EQUIPMENT BAY FWD RACK
Figure 2: Integrated Air System Controller Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_IASC_ComponentLocation.fm
21 - Environmental Control 21-90 Integrated Air System Controller
DETAILED DESCRIPTION The IASC is a fully redundant controller. In case of failure of one channel, the other channel can control the ECS without loss of any function. The AIR SYSTEM FAULT advisory message is displayed for the failure of a channel. The IASC safety channel does an additional monitoring of all the most critical functions of the air systems. In the cooling and temperature control system, it monitors for overheat on temperature sensors and reports to the data concentrator unit module cabinet (DMC). In the avionics cooling and extraction system, it controls the extraction fans (EFAN) for the air extraction system, and reports to the DMC.
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CS130_21_L3_IASC_DetailedDescription.fm
21 - Environmental Control 21-90 Integrated Air System Controller
LEGEND IASC 1
AEV CSOV DMC
Channel A
Channel B
Safety Channel C
• L pack discharge disconnect and monitoring • Aft CSOV control and monitoring • Fwd galley heater temperature monitoring • EFAN no. 1 speed monitoring and control on OMS • Mid equipment bay temp monitoring • Control of both fwd/mid AEV and bufan • Fwd inlet fan control and monitoring
• L pack discharge disconnect and monitoring • R fan control and monitoring • Aft CSOV control and monitoring • Fwd galley heater temperature monitoring • Fwd equipment bay temperature monitoring • Fwd inlet fan control and monitoring
• EFAN no. 1 operation selection and failure detection
Data Exchange
EFAN IASC OMS
Avionics Exhaust Valve Cargo Shutoff Valve Data Concentrator Module Cabinet Exhaust Fan Integrated Air System Controller Onboard Maintenance System ARINC 429
DMC
IASC 2
Channel B
Safety Channel C
• R pack discharge disconnect and monitoring • Fwd CSOV control and monitoring • Aft galley heater temperature monitoring • EFAN no. 2 speed monitoring and monitoring report on OMS • Fwd equipment bay temperature monitoring • Control of both fwd/mid AEV and bufan • Fwd inlet fan control and monitoring • Mid ground valve control and monitoring
• R pack discharge disconnect and monitoring • R fan control and monitoring • Fwd CSOV control and monitoring • Aft galley heater temperature monitoring • Fwd equipment bay temperature monitoring • Mid inlet fan control and monitoring • Mid ground valve control and monitoring
• EFAN no. 2 operation selection and failure detection
CS1_CS3_2190_006
Channel A
Figure 3: Integrated Air System Controller Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-90 Integrated Air System Controller
MONITORING AND TESTS The following page provides the INFO messages associated with the integrated air system controller:
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CS130_21_L2_IASC_MonitoringTests.fm
21 - Environmental Control 21-90 Integrated Air System Controller
INFO MESSAGES Table 1: INFO Messages MESSAGE
LOGIC
21 AIR SYSTEM FAULT - IASC 1A INOP
No data received from IASC1A confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 1B INOP
No data received from IASC1B confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 2A INOP
No data received from IASC2B confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 2B INOP
No data received from IASC2B confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 1C INOP
No data received from IASC1C confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 2C INOP
No data received from IASC2C confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - L IASC ARINC INPUT LOSS
No data from DMC1 or from DMC2 received by IASC1.
21 AIR SYSTEM FAULT - R IASC ARINC INPUT LOSS
No data from DMC1 or from DMC2 received by IASC2.
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21 - Environmental Control 21-90 Integrated Air System Controller
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM TEST The electrical environmental system components can be tested using the IBIT-ECS test on the onboard maintenance system (OMS). Follow the preconditions before starting the test.
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CS130_21_L2_IASC_PracticalAspects.fm
21 - Environmental Control 21-90 Integrated Air System Controller
MAINT MENU
RETURN TO FAULT MSGS
IASC1 chB-IBIT ECS
001 /004 IBIT ECS
EQUIPMENT TESTED: FCV, TASOV, TAPRV, TAV, DTS, VENTS, TCV, PACKS SENSORS, RARV
PRECONDITIONS
1 2 3 4 5
-
MAKE MAKE MAKE MAKE MAKE
SURE SURE SURE SURE SURE
WEIGHT ON WHEEL L BLEED SW IS SELECTED OFF R BLEED SW IS SELECTED OFF ALL ENGINE ARE OFF APU BLEED SW IS SELECTED OFF TEST INHIBIT/INTERLOCK CONDITIONS
WEIGHT ON WHEEL LEFT BLEED SWITCH
YYY
RIGHT BLEED SWITCH
YYY
ENGINE STATUS
YYY
APU BLEED SWITCH
YYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2190_003
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 4: Environmental Control System OMS Page Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-11
CS130_21_L2_IASC_PracticalAspects.fm
21 - Environmental Control 21-53 Flow Control System
21-53 FLOW CONTROL SYSTEM GENERAL DESCRIPTION The flow control system controls the airflow, airflow rate, and the temperature of bleed air provided into each air conditioning unit. The flow control system is installed on the system feed line, upstream of the air conditioning units. A trim air system is supplied with bleed air downstream of the flow control system. There is one system associated with each of the air conditioning units, installed on the aircraft. Each flow control system provides: •
Measurement and control of the air flow supplied to the air conditioning packs
•
Pack shutoff control
•
Pack overheat protection
An optional ozone converter may be installed in the bleed air ducting, upstream of the flow sensing venturi. The ozone converters convert ozone concentrations into oxygen, which prevents high ozone concentrations in the cabin. This improves the quality of air for high altitude or long duration flights. A second option adds a combined ozone and volatile organic compounds (VOC) converter.
The FCV is powered and controlled by the IASC. The left IASC controls the left FCV and the right IASC controls the right FCV. The IASC positions the FCV butterfly, using a torque motor, to limit the pressure and temperature of the air entering the pack. The FCV is spring-loaded closed. In the event of a power loss, the valve closes. A minimum 15 psi pressure is required to drive the valve in the open position. The maximum regulated pressure is limited to 43.5 +/- 3 psi. Each FCV is controlled by the associated PACK PBA on the AIR panel in the flight deck. The FCV is normally automatically controlled by the IASC but can be selected closed the L or R PACK PBA. During flight with reduced passenger loads, the system reduces the flow schedule in order to reduce bleed air consumption. Using the FMS data to capture passenger load, the IASC reduces the flow schedule based on the actual quantity of passengers onboard. The PACK FLOW PBA on the AIR panel, can be used to adjust the flow control to the high flow mode, which is used for smoke and odor clearance as well as increased pack performance.
The mass airflow into the air conditioning pack is measured by a flow sensor venturi (FSV) and associated sensors. The FSV is used to calibrate and measure the flow into the pack and trim air system. A pack inlet flow sensor (PIFS) provides differential pressure measurement across the flow sensor venturi to the IASC for mass flow computation and flow control valve (FCV) control. The pack inlet pressure sensor (PIPS) provides pack inlet static pressure to the IASC for mass flow computation and bleed air manifold pressure monitoring. The bleed air manifold pressure is displayed on the AIR synoptic page.
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21-12
CS130_21_L2_FCS_General.fm
21 - Environmental Control 21-53 Flow Control System
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
OFF
45
OPEN
XBLEED
PSI
45 PSI
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL
FMS
FAIL
SYNOPTIC PAGE – AIR
OFF R BLEED
DMC 1 DMC 2
FAIL
APU BLEED
IASC 1
OFF
FAIL
VOC
Volatile Organic Compounds
T/M
FSV
FCV
LEFT AIR CONDITIONING PACK
To Trim Air System
CS1_CS3_2150_013
DMC FCV FMS FSV IASC PIFS PIPS
ARINC 429 Data Concentrator Unit Module Cabinet Flow Control Valve Flight Management System Flow Sensor Venturi Integrated Air System Controller Pack Inlet Flow Sensor Pack Inlet Pressure Sensor
OZONE AND VOC CONVERTER (OPTIONAL)
L PIFS 2
LEGEND
Hot Air From Bleed Air System
L PIPS 2
AIR PANEL
L PIPS 1
L PIFS 1
OFF
Figure 5: Flow Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-13
CS130_21_L2_FCS_General.fm
21 - Environmental Control 21-53 Flow Control System
COMPONENT LOCATION The flow control system has the following components installed in the wing-to-body fairing (WTBF): •
Flow sensor venturi (FSV)
•
Flow control valve (FCV)
•
Ozone converter or combined ozone and volatile organic compounds (VOC) converter (Optional)
•
Pack inlet pressure sensor (PIPS) (Refer to figure 7)
•
Pack inlet flow sensor (PIFS) (Refer to figure 7)
FLOW SENSOR VENTURI The flow sensor venturi is located in the WTBF. FLOW CONTROL VALVE The flow control valve is located in the WTBF. OZONE CONVERTER/OZONE AND VOC CONVERTER (OPTIONAL) The ozone converter is installed upstream of the flow sensor venturi.
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21-14
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
Right Ozone and VOC Converter
Right Flow Sensor Venturi
Right Flow Control Valve
Left Ozone and VOC Converter
LEGEND VOC
Left Flow Control Valve
Volatile Organic Compounds
CS1_CS3_2150_030
Left Flow Sensor Venturi
Figure 6: Flow Sensor Venturi, Flow Control Valve, and Ozone Converter Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-15
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
PACK INLET PRESSURE SENSORS The pack inlet pressure sensors (PIPSs) are mounted on the pressure bulkhead in the wing-to-body fairing (WTBF). PACK INLET FLOW SENSORS The pack inlet flow sensors (PIFSs) are mounted on the pressure bulkhead in the WTBF.
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21-16
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
Right PIF 1 Sensor
Right PIF 2 Sensor
Right PIP 2 Sensor
Right PIP 1 Sensor
Left PIP 2 Sensor Left PIP 1 Sensor
Left PIF 1 Sensor
CS1_CS3_2150_005 50 005
Left PIF 2 Sensor
NOTE PIP/PIF SENSORS
Left side shown.
Figure 7: PIP Sensor and PIF Sensor Locations Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-17
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
CONTROLS AND INDICATIONS CONTROLS The L and R PACK PBA are used to select the flow control valves closed. An OFF legend on the PACK PBA indicates the valve is closed. The amber FAIL light indicates a pack fault. The HI FLOW PBA is used to increase the pack flow manually.
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21-18
CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
AIR COCKPIT
C
CABIN
H
C
H
FWD
MAN TEMP
C
CARGO
VENT
OFF
LO HEAT
OVERHEAD PANEL
AFT
OFF
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
VENT
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
H
HI HEAT
ON
FAIL
AFT
FAIL OFF R BLEED FAIL
APU BLEED
OFF
OFF
AIR PANEL
CS1_CS3_2160_006
FAIL
Figure 8: Flow Control Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-19
CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
INDICATIONS A HI indication appears in the pack symbol on the AIR synoptic page indicating the packs are in HI FLOW mode. The bleed air manifold pressure is supplied by the PIPS.
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CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
STATUS
AIR IR
DOOR OR
ELECC
FLT CTR CTRL TRLL
FUEL FUE EL
HYD YD
AVIONIC IONIC
INFO
CB
FWD
CKPT 23 °C 22 °C
CARGO
LO
15°CC
10 °C °
AFT
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2 15 °CC
VALVE POSITION NORMAL OPERATION Symbol
15 °CC
FLOW LINES Symbol
Condition
Condition Normal flow Closed
L PACK
HI
No flow
RAM AIR
TRIM AIR
Open with flow
R PACK
HI
Abnormal flow (overheat, out of range pressure, leak)
Open with no flow
VALVE POSITION FAIL OPERATION Symbol
45 PSI
XBLEED
45 PSI
Condition Closed Open with flow Open with no flow
CS1_CS3_2120_007
Invalid
APU
SYNOPTIC PAGE – AIR
Figure 9: Flow Control Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-21
CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
DETAILED DESCRIPTION The main operating modes for the system are the following:
FLOW CONTROL NORMAL OPERATION The flow is automatically controlled by the integrated air system controller (IASC) depending on the system configuration according to preset flow schedules. The FCV opens when the IASC provides a signal to the torque motor on the FCV and there is a minimum of 15 psi available from the bleed air system. The IASCs use the actual flow measured at the flow sensing venturi and the bleed temperature sensor (BTS) to modulate the FCV to obtain the required flow schedule. During flight with reduced passenger loads, the system automatically reduces the flow schedule in order to reduce bleed air consumption. Using the flight management system (FMS) passenger load data, the IASC reduces the flow schedule based on the actual quantity of passengers onboard. In the auto mode, the bleed air system uses auxiliary power unit (APU) bleed on the ground and during takeoff. The system switches to engine bleed air based on aircraft configuration. For additional information, refer to ATA 36 - Pneumatic.
•
Two bleed sources dual-pack operation
•
Two bleed sources single-pack operation
•
One bleed source dual-pack operation
•
Trim air only during emergency ram air operation
Pull-Up and Pull-Down Sequence When the APU is supplying bleed air on the ground and the actual zone temperatures vary by more than 10°C from the required temperature, the system automatically goes into pull-up or pull-down mode. Pull-up or pull-down modes use an increased flow schedule. During pull-up mode, additional heating is supplied by the trim air system to improve the pull-up efficiency.
On the ground, in dual pack or single pack operation, the APU flow demand is set to 100%, and the flow control logic uses a flow schedule based on: •
Aircraft version (CS100 or CS300)
•
Altitude (ranging from -2,000 ft to 14,500 ft)
•
Outside air temperature (-50°C to 50°C) (-58°F to 122°F)
In flight, the APU flow demand is set to 100%, and the system regulates the FCV according to the engine flow schedule.
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21-22
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
From L Bleed Air System
L PIPS 1
L PIFS 1
Recirc Air
L BTS
L FCV T/M
OZC
L PACK
IASC 1
APU
M To Trim Air System
DMC
IASC 2
R PIPS 1
R PIFS 1
APU ECU
MIX MANIFOLD
LEGEND
T/M
R PACK
OZC R PIPS 2
From R Bleed Air System
FSV R PIFS 2
R BTS
R FCV
Recirc Air
APU BTS DMC ECU FCV FSV HPGC IASC OZC PIFS PIPS
Hot Air Cold Air Auxiliary Power Unit Bleed Temperature Sensor Data Concentrator Unit Module Cabinet Electronic Control Unit Flow Control Valve Flow Sensor Venturi High-Pressure Ground Connection Integrated Air System Controller Ozone Converter Pack Inlet Flow Sensor Pack Inlet Pressure Sensor ARINC 429
CS1_CS3_2150_010
L PIFS 2
HPGC
L PIPS 2
FSV
Figure 10: Flow Control Normal Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-23
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
FLOW CONTROL To perform the flow control, the IASCs use the actual flow measured by: •
Pack inlet pressure sensor (PIPS)
•
Pack inlet flow sensor (PIFS)
•
Bleed temperature sensor (BTS)
The flow reference is function of: •
Number of passengers
•
Number of bleed air users
•
Aircraft altitude
•
Compressor discharge temperature sensor value
•
Bleed source selection
The flow requirements take into account the total passenger loads as well as the three cockpit members and five flight attendants.
NOTE 1. In dual-pack operation, with FWD cargo heat in either the LO or HI setting, and the recirculation air on, the flow schedule on the right side is increased to account for the added FWD cargo trim air. When the recirculation air system is off, but the FWD cargo heating is on, then both left and right FCVs are supplied with the same flow. 2. During single pack operation, the aircraft is altitude limited. The loss of a PIPS, PIFS, or an IASC control channel results in a PACK FAULT advisory message. Normal operation continues with the remaining sensors or the working channel. The loss of both PIPs, PIFs, or IASC control channels for either FCV is indicated by a L or R PACK FAIL caution message and an L(R) PACK PBA FAIL light. The pack must be shutdown by selecting the PACK PBA to OFF. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-24
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
IASC 1 Aircraft Altitude Number of Bleed Users Number of Passengers
FLOW CONTROL VALVE Flow Scheduled Value
T/M
CDTS Value Bleed Source Selection + -
L PIPS 1 L PIPS 2 Flow Measurement L PIFS 1 L PIFS 2
BTS
LEGEND Bleed Air Bleed Temperature Sensor Compressor Discharge Temperature Sensor Integrated Air Systems Controller Pack Inlet Flow Sensor Pack Inlet Pressure Sensor
CS1_CS3_2150_011
BTS CDTS IASC PIFS PIPS
Figure 11: Flow Control Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-25
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
FLOW CONTROL MODES Flow Priority When the flow measured is less than 80% of the minimum required flow for 3 minutes in single pack operation, or 90% in dual-pack operation, the flow priority mode is enabled. The flow priority circuit monitors for a low flow supply condition based on operating mode and passenger load as input in the FMS. The corresponding pack temperature control valve (TCV) is commanded open by the IASC. This allows increased air flow either until the flow target is achieved, or until the flight deck and cabin zone temperatures reach 35°C (95°F). The ram air regulating valve (RARV) opens to maintain the pack temperature. In this mode of operation, the flight deck and cabin zone temperature control are not available as the pack outlet temperature may be higher than the selected value. If the IASC detects a low flow condition for at least 3 minutes, an engine indication and crew alerting system (EICAS) L or R PACK FAIL advisory message is displayed.
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21-26
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control
L PIPS 1
L PIFS 1
21-53 Flow Control System
L FCV
CDTS
OZC
Reheater T/M
TURB
COMP
PTS
Condenser PDTS
L PIPS 2
L PIFS 2
FSV
M
M
PDPS
Air Cycle Machine TCV AIR CONDITIONING PACK
RARV
IASC 1
FWD CABIN VENTS
IASC 2
LEGEND Data Concentrator Unit Module Cabinet Flow Control Valve Flight Management System Integrated Air System Controller Ozone Converter Pack Inlet Flow Sensor Pack Inlet Pressure Sensor Pack Temperature Sensor Ram Air Regulating Valve Temperature Control Valve Ventilated Temperature Sensor ARINC 429
DMC 1 and DMC 2
AFT CABIN VENTS
FLIGHT DECK VENTS
FMS
CS1_CS3_2150_018
DMC FCV FMS IASC OZC PIFS PIPS PTS RARV TCV VENTS
Figure 12: Flow Control Priority Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
Hi Flow
FLOW CONTROL SHUTOFF
Whenever the PACK FLOW PBA is selected on the AIR panel, the flow control system switches to the high flow mode. The high flow mode corresponds to the maximum allowable flow for the system and it overrides all other flow control schedules. This mode is used to help evacuate cabin odors, smoke, or to increase the system air conditioning performance. The high flow mode can exchange the cabin air in less than 5 minutes in either dual and single PACK operating modes.
The flow control valve is closed when: •
The associated PACK PBA is selected OFF
•
Engine start in flight
•
Engine start on ground plus 30 seconds
•
Ditching PBA pressed
The HI flow mode is also automatically enabled when the auxiliary power unit (APU) is operating on the ground or when the required mix manifold temperatures decrease below 12°C (54°F) or increase above 60°C (140°F).
•
Bleed leak in pack zone as reported by the bleed air leak detection and overheat system
In flight, if the EMER DEPRESS PBA is selected, or if the cabin altitude increases above 15,000 ft, the flow control automatically switches to high flow mode.
When the FCV is commanded open with maximum torque motor current and the PIPS pressure is at least 15 psi and a low flow has been detected, a L or R PACK FAIL caution message is displayed.
FLOW CONTROL VALVE FAILURE
The FCV is detected failed open when it is commanded closed and the corresponding flow computed by the IASC is greater than 30 lb/min for more than 30 seconds. The L or R PACK FAIL caution message is displayed. This message is inhibited during emergency ram air mode with trim air ON since the flow through the FSV includes both the FCV and trim air system flow, which can exceed 30 lb/min in this mode.
NOTE During single engine bleed with: Wing anti-ice on (WAIS) Approach Steep approach (CS100 only) The flow control is limited to the reference flow, and HI flow is not available.
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CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control
L FCV
CDTS
OZC
Reheater
EEC FCV FSV IASC PDPS PDTS PIFS PIPS PTS RARV TCV
TURB
COMP
PDTS PDPS
M
TCV
AIR CONDITIONING PACK Loop A
Loop B
DMC
M Bleed Air Warm Air Cold Air RARV Water Air Cycle Machine Compressor Discharge Temperature Sensor Data Concentrator Unit Module Cabinet Electronic Engine Control Flow Control Valve Flow Sensor Venturi Integrated Air System Controller Pack Discharge Pressure Sensor Pack Discharge Temperature Sensor Pack Inlet Flow Sensor Pack Inlet Pressure Sensor Pack Temperature Sensor Ram Air Regulating Valve Temperature Control Valve ARINC 429 Bleed Air Detection Loop A Bleed Air Detection Loop B
Condenser
Air Cycle Machine
LEGEND
ACM CDTS
PTS
T/M
L PIPS 2
L PIFS 2
FSV
IASC 1
IASC 2
PRESSURIZATION MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK
AUTO PRESS
L BLEED
FAIL
FAIL
ON
MAN CLSD
OFF
DITCHING
L EEC and R EEC
PRESSURIZATION PANEL
OFF
R PACK XBLEED
FAIL
MAN
OPEN
DMC AUTO
MAN OPEN
FAIL OFF R BLEED
APU BLEED FAIL OFF
AIR PANEL
FAIL OFF
CS1_CS3_2150_012
L PIPS 1
L PIFS 1
21-53 Flow Control System
Figure 13: Flow Control Shutoff Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-29
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the flow control system:
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21-30
CS130_21_L2_FCS_MonitoringTests.fm
21 - Environmental Control 21-53 Flow Control System
CAS MESSAGES Table 5: INFO Messages
Table 2: CAUTION Message MESSAGE L/R PACK FAIL
LOGIC
MESSAGE
Left/right pack inoperative (loss of control or monitoring [CDTS or PDTS failure] or pack low flow detected).
Table 3: ADVISORY Message MESSAGE PACK FAULT
LOGIC Loss of redundant or non-critical function for the air conditioning system.
Table 4: STATUS Message MESSAGE PACK FLOW HI
LOGIC CABIN AIR FLOW selected HI and confirmed HI by the controllers.
LOGIC
21 L PACK FAIL L FLOW CTRL VLV INOP
Left FCV failed open.
21 R PACK FAIL R FLOW CTRL VLV INOP
Right FCV failed open.
21 AIR SYSTEM FAULT - L PACK PRESS SNSR REDUND LOSS
Left pack pressure sensor redundant loss (PIFS or PIPS).
21 AIR SYSTEM FAULT - R PACK PRESS SNSR REDUND LOSS
Right pack pressure sensor redundant loss (PIFS or PIPS).
36 L BLEED FAIL - Left PIPS 1 and PIPS 2 loss. L PACK INLET PRESS SNSR INOP 36 R BLEED FAIL - R PACK INLET PRESS SNSR INOP
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TECHNICAL TRAINING MANUAL
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Right PIPS 1 and PIPS 2 loss.
21-31
CS130_21_L2_FCS_MonitoringTests.fm
21 - Environmental Control 21-53 Flow Control System
PRACTICAL ASPECTS The flow control valve can be dispatched in the closed position by capping and stowing the electrical connector and then removing the deactivation screw and installing it in the manual position indicator. If the valve is open, the manual position indicator can be used to close the valve using a wrench.
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21-32
CS130_21_L2_FCS_PracticalAspects.fm
21 - Environmental Control 21-53 Flow Control System
Visual Indicator and Lockout Mechanism
Deactivation Screw Position for Valve Lockout
CS1_CS3_2150_006 006
Deactivation Screw
Figure 14: Flow Control Valve Lockout Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_FCS_PracticalAspects.fm
21 - Environmental Control 21-51 Cooling
21-51 COOLING
ACM compresses and heats the air. The air passes through the main heat exchanger for further cooling before entering the reheater. The reheater heats the air entering the turbine inlet, where the remaining water is vaporized in order to prevent damage to the ACM.
GENERAL DESCRIPTION Cooling air is provided from two sources: •
Air conditioning units
•
Low-pressure ground cart
The air conditioning units, or packs are available on the ground or in flight. Each pack consists of: •
Air cycle machine
•
Dual-heat exchanger
•
High-pressure water extraction system
•
Temperature control valve
Air exiting the water extractor enters the turbine inlet where it expands and cools. This air is fed to the mix manifold for distribution to the aircraft.
The air cycle machine (ACM) consists of a turbine, compressor, and a fan connected by a common shaft. The ACM shaft is supported by air bearings. The dual-heat exchanger consists of a primary heat exchanger used to cool air from the bleed air system, and a main heat exchanger used to cool air from the air cycle machine (ACM) compressor discharge. The high-pressure water extraction system consists of a reheater, condenser, and a water separator. The extracted water is sprayed across the heat exchanger inlet to improve the cooling performance. The temperature control valve (TCV) bypasses bleed air around the ACM and exhausts it at the turbine outlet. This controls the pack outlet temperature and prevents pack icing. Each pack is controlled by an integrated air system controller (IASC). IASC 1 controls the left pack, and IASC 2 controls the right pack. Bleed air enters the pack through the flow control system. The bleed air is cooled by the primary heat exchanger before entering the ACM. The Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The air in the reheater starts cooling and any water vapor starts condensing. This air enters the condenser where it is cooled by the cold turbine outlet air. The water and air pass through the water extractor where the water is slung against the walls and is discharged through a line that runs to a spray tube upstream of the dual-heat exchanger. The water sprays across the face of the heat exchanger, increasing its cooling performance.
The water extraction line is monitored by the pack temperature sensor (PTS). If the air is too cold, there is a possibility of ice forming and damaging the ACM. To prevent this, the temperature control valve (TCV) is opened to allow some hot air to bypass the ACM, raising the temperature of the air. The dual-heat exchanger is cooled by ram air in flight. On the ground, the ACM fan draws air across the heat exchanger when the ACM is operating. The amount of cooling air drawn across the heat exchanger depends on the position of the ram air regulating valve (RARV). The pack temperature is controlled by modulating the RARV and the TCV. Pack overheat protection is provided by the compressor discharge temperature sensor and the pack discharge temperature sensor. If either sensor detects an overheat condition, a caution message is displayed and the pack is turned off using the PACK PBA which closes the flow control valve. On the ground, the low-pressure ground connection allows the use of an external ground cart to cool the aircraft. The low-pressure ground connection allows conditioned air to enter the left pack discharge duct downstream of the pack.
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CS130_21_L2_Cooling_General.fm
21 - Environmental Control 21-51 Cooling
From Bleed Air System
Ram Air Inlet
To Trim Air System Flow Control Valve
T/M
CDTS PTS
MIXTS
TURB
COMP
Reheater
PDTS
Condenser
To Mix Manifold
Air Cycle Machine M
M Temperature Control Valve
Ram Air Outlet
Water Extractor PDPS
Ram Air Regulating Valve
LEGEND
CS1_CS3_2160_003
CDTS MIXTS PDPS PDTS PTS
Bleed Air Warm Air Cold Air Water Compressor Discharge Temperature Sensor Mix Manifold Temperature Sensor Pack Discharge Pressure Sensor Pack Discharge Temperature Sensor Pack Temperature Sensor
Figure 15: Air Conditioning System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-35
CS130_21_L2_Cooling_General.fm
21 - Environmental Control 21-51 Cooling
COMPONENT LOCATION The following cooling system components are located in the wing-to-body fairing (WTBF): •
Air cycle machine
•
Dual-heat exchanger
•
Reheater condenser
•
Temperature control valve
•
Compressor discharge temperature sensor (CDTS)
•
Pack temperature sensor (PTS)
•
Plenum
•
Pack discharge temperature sensor (PDTS)
•
Ram air regulating valve (RARV)
•
RAM air duct
•
Water sprayer (Refer to figure 17)
•
Water extractor (Refer to figure 17)
The following cooling system component is located in the environmental control system bay: •
Pack discharge pressure sensor (Refer to figure 18)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
Pack Discharge Temperature Sensor
Water Extractor
Pack Temperature Sensor
Condenser/Reheater
Compressor Discharge Temperature Sensor Air Cycle Machine
Temperature Control Valve Plenum Dual Heat Exchanger
CS1_CS3_2150_019 S1 CS3 2150 019
Ram Air Duct
Figure 16: Air Conditioning Components Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-37
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
WATER EXTRACTOR The water extractor is mounted in front of the condenser/reheater. WATER SPRAYER The water sprayer is located in the ram air duct forward of the dual-heat exchanger. The sprayer discharge is oriented forward to optimize distribution across the face of the heat exchanger.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-38
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
Water Sprayer
Heat Exchanger Inlet
Water Extractor Water Extractor Drain
CS1_CS3_2150_007
Condenser/Reheater
Figure 17: Water Extractor and Water Sprayer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-39
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
PACK DISCHARGE PRESSURE SENSOR The pack discharge pressure sensor is located in the environmental control system bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-40
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
Pack Discharge Pressure Sensor Line
CS1_CS3_2150_020
Pack Discharge Pressure Sensor
Figure 18: Pack Discharge Pressure Sensor Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-41
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
COMPONENT INFORMATION TEMPERATURE CONTROL VALVE The temperature control valve (TCV) controls the pack outlet temperature, by extracting hot air downstream from the primary heat exchanger and reinjecting it at the turbine. The TCV is driven by a stepper motor. Microswitches provide position feedback to the IASC. In case of a power loss, the TCV remains in the last position. The TCV has a manual override lever to open or close the valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-42
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
Position Indicator
CS1_CS3_2160_013
Manual Override Lever
Figure 19: Temperature Control Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-43
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
RAM AIR REGULATING VALVE The ram air regulating valve (RARV) is used to regulate the ram air flow used on cold side of the PACK dual-heat exchanger. The RARV is driven by a stepper motor. Microswitches provide open and close position feedback. There is always flow through the valve, even when it is closed due to the fact that the butterfly valve is smaller than the valve housing. In case of power loss on the actuator, the RARV remains in its last position. The RARV has a manual override to open the valve for dispatch.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-44
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
Manual Lever
CS1_CS3_2150_027
Ram Air Regulating Valve
Figure 20: Ram Air Regulating Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-45
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
CONTROLS AND INDICATIONS CONTROLS The PACK PBA turns off the pack. An OFF legend on the PACK PBA indicates the pack is off. The PACK PBA FAIL legend indicates a fault has occurred that requires the pack to be shutdown.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-46
CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
AIR COCKPIT
C
FWD
H
CABIN
C
AFT
H
FWD
MAN TEMP
CARGO
VENT
OFF
C
LO HEAT
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK
OPEN
R PACK XBLEED
FAIL
MAN CLSD
OFF
AUTO
MAN OPEN
L BLEED
OFF
OFF R BLEED FAIL
APU BLEED
OFF
FAIL OFF
AIR PANEL
CS1_CS3_2150_025
FAIL
FAIL
Figure 21: Air Conditioning Pack Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-47
CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
INDICATIONS The pack status is shown on the AIR synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°C
10 °C
15 °C
15 °C RAM AIR
TRIM AIR
PACKS Symbol
R PACK
45 PSI
XBLEED
45
Condition
L PACK
Fail or leaked
L PACK
Active
L PACK
Invalid
L PACK
Normal OFF
PSI
CS1_CS3_2150_026
L PACK
APU
SYNOPTIC PAGE – AIR
Figure 22: Air Conditioning Pack Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-49
CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
DETAILED DESCRIPTION PACK TEMPERATURE REGULATION The pack temperature control is performed by two regulation loops within the IASC. The system controls the RARV and TCV simultaneously. The TCV is kept as close to the fully closed position as possible, minimizing the ram air flow, and the resulting induced drag penalty. On the ground, with the PACK selected ON, the RARV is commanded fully open above -19°C. Below this temperature the normal RARV/TCV control applies. In flight, when the PACK is OFF, the RARV is in its closed position. When the PACK is selected ON, and the flaps or the landing gear are not fully retracted, the RARV is commanded fully open. The TCV regulation loop compares mix manifold temperature sensor (MIXTS) values or PDTS if MIXTS is not available to a pack reference temperature. The RARV regulation loop compares compressor discharge temperature sensor (CDTS) value to a CDTS reference function of pack reference temperature. When the TCV or RARV fails, pack temperature control is maintained with the remaining TCV or RARV. An L or R PACK FAIL advisory message is displayed on EICAS for a TCV or RARV failure. PACK ICING PROTECTION To prevent ice buildup in the pack discharge outlet, the IASC opens the temperature control valve (TCV) to melt the ice. Icing protection is only available below 30,000 ft when the MIXTS is below 15°C (59°F). The ice formation is detected by the pack discharge pressure sensor (PDPS). The PDPS senses the difference between the pack discharge outlet and the cabin pressure .
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The TCV provides de-icing when the PDPS senses the differential pressure in the distribution duct is greater than 1.59 psi for more than 5 seconds. This mode is deactivated when the PDPS is less than 0.72 psi or after 30 seconds. In the event of a total loss of the PDPS, the pack discharge temperature is limited to 5°C (41°F) to prevent the formation of ice. PACK OVERHEAT PROTECTION The IASCs monitor the packs for an overheat condition using a compressor discharge temperature sensor (CDTS). The CDTS supplies an output to each IASC. The IACS safety channels monitor the CDTS and if the temperature exceeds 232°C (450°F) for 30 seconds or 260°C (500°F) for 5 seconds, an L or R PACK OVHT caution message is displayed. The pack must be shutdown manually using the PACK PBA. Selecting the PACK PBA FAIL light to OFF closes the flow control valve (FCV). To prevent a pack overheat condition, the system reduces the FCV flow as soon as the CDTS exceeds 220°C (428°F) and maintains the reduced flow until the temperature drops below this value. If the CDTS fails completely, the pack has no overheat protection and the same L or R PACK FAIL caution message is displayed. PACK DISCHARGE OVERHEAT PROTECTION In the normal mode, the pack discharge temperature is limited to 70°C (158°F). The pack discharge is monitored by the pack discharge temperature sensor (PDTS). An overheat occurs when the pack discharge temperature exceeds 85°C (185°F) for 30 seconds. When IASC safety channel detects a pack discharge overheat, a L or R PACK OVHT caution message is displayed. The FCV must be closed by selecting the PACK PBA FAIL light to OFF. If the PDTS fails completely, pack discharge overheat protection is not available, a L or R PACK FAIL caution message is displayed.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-50
CS130_21_L3_Cooling_DetailedDescription.fm
21 - Environmental Control 21-51 Cooling
LEGEND Ram Air Inlet
FCV CDTS
TURB
Ram Air Regulating Valve
PTS
Air Cycle Machine
M
M
To Mix Manifold
Condenser PDTS
COMP
Reheater
MIXTS
T/M
From Bleed Air System
CDTS DMC FCV MIXTS PDPS PDTS PTS TCV
Bleed Air Cold Air Warm Air Water Compressor Discharge Temperature Sensor Data Concentrator Unit Module Cabinet Flow Control Valve Mix Manifold Temperature Sensor Pack Discharge Pressure Sensor Pack Discharge Temperature Sensor Pack Temperature Sensor Temperature Control Valve ARINC 429
TCV
Channel B
AIR COCKPIT
C
FWD
H
C
OFF
CABIN
H
FWD
MAN TEMP
VENT
IASC 1 Channel B
AFT
C
CARGO LO HEAT
Channel A
H
AFT
OFF
VENT
Safety Channel
DMC 1 and DMC 2
Channel A
Safety Channel
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
OPEN
CS1_CS3_2150_015
IASC 2
PDPS
Pressurized Cabin Underfloor
Figure 23: Air Conditioning System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-51
CS130_21_L3_Cooling_DetailedDescription.fm
21 - Environmental Control 21-51 Cooling
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the air conditioning system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-52
CS130_21_L2_Cooling_MonitoringTests.fm
21 - Environmental Control 21-51 Cooling
CAS MESSAGES
Table 9: INFO Messages Table 6: CAUTION Messages
MESSAGE L/R PACK FAIL
MESSAGE
LOGIC Left pack inoperative (loss of control or monitoring). Loss of L PACK leak detection capability. Loss of both IASC1 channels A and B.
L/R PACK FAIL
Right pack inoperative (loss of control or monitoring). Loss of R PACK leak detection capability. Loss of both IASC2 channels A and B.
L PACK OVHT
Left compressor discharge overheat or left pack discharge overheat.
R PACK OVHT
Right compressor discharge overheat or left pack discharge overheat.
Table 7: ADVISORY Message MESSAGE PACK FAULT
LOGIC Loss of redundant or non-critical function for the air conditioning system.
Table 8: STATUS Message MESSAGE L/R PACK OFF
LOGIC Left/right pack selected OFF and confirmed OFF.
Table 9: INFO Messages MESSAGE 21 PACK FAULT MIX MANF TEMP SNSR TOTAL LOSS
LOGIC MMTS from both packs are lost.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
21 PACK FAULT MIX MANF TEMP SNSR REDUND LOSS
Up to 3 MMTS sensing elements redundant loss or any sensing element drift.
21 PACK FAULT L BYPASS VLV INOP
Left TCV failed in position or position unknown.
21 PACK FAULT R BYPASS VLV INOP
Right TCV failed in position or position unknown.
21 L PACK FAIL L FLOW CTRL VLV INOP
Left FCV failed open.
21 R PACK FAIL R FLOW CTRL VLV INOP
Right FCV failed open.
21 L PACK FAIL L PACK INOP
L Pack failed for any reason except FCV failed open case.
21 R PACK FAIL R PACK INOP
R Pack failed for any reason except FCV failed open case.
21 PACK FAULT L PACK TEMP SNSR REDUND LOSS
Left CDTS or PDTS drift or single-channel failure.
21 PACK FAULT R PACK TEMP SNSR REDUND LOSS
Right CDTS or PDTS drift or single-channel failure.
21 PACK FAULT L PACK DISCH PRESS SNSR INOP
Left PDPS loss.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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21 - Environmental Control 21-51 Cooling
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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21 - Environmental Control 21-51 Cooling
Table 9: INFO Messages MESSAGE 21 PACK FAULT R PACK DISCH PRESS SNSR INOP
Table 9: INFO Messages
LOGIC
MESSAGE
Right PDPS loss.
21 L PACK OVHT - L PACK INOP
LOGIC L pack overheat triggered for any reason except RARV failed in position case.
21 R PACK OVHT R pack overheat triggered for any reason except RARV - R PACK INOP failed in position case.
36 L BLEED FAIL - Left PIPS 1 and PIPS 2 loss. L PACK INLET PRESS SNSR INOP
21 R PACK OVHT Left PTS drift or both channels failure. - L PACK TEMP SNSR INOP
36 R BLEED FAIL - R PACK INLET PRESS SNSR INOP
Right PIPS 1 and PIPS 2 loss.
21 R PACK OVHT Right PTS drift or both channels failure. - R PACK TEMP SNSR INOP L pack discharge duct disconnected.
21 AIR SYSTEM FAULT - L PACK PRESS SNSR REDUND LOSS
Left pack pressure sensor redundant loss (PIFS or PIPS).
21 L PACK FAIL L PACK DISCHARGE DUCT DISCONNECT
21 AIR SYSTEM FAULT - R PACK PRESS SNSR REDUND LOSS
Right pack pressure sensor redundant loss (PIFS or PIPS).
21 R PACK FAIL R PACK DISCHARGE DUCT DISCONNECT
R pack discharge duct disconnected.
21 PACK FAULT L RAM AIR REG VLV INOP
Left RARV failed in position or unknown position.
21 PACK FAULT R RAM AIR REG VLV INOP
Right RARV failed in position or unknown position.
21 L PACK OVHT Left RARV failed in position leading to pack overheat. - L RAM AIR REG VLV INOP 21 R PACK OVHT Right RARV failed in position leading to pack overheat. - R RAM AIR REG VLV INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-55
CS130_21_L2_Cooling_MonitoringTests.fm
21 - Environmental Control 21-52 Ram Air System
21-52 RAM AIR SYSTEM GENERAL DESCRIPTION In case of dual pack loss, the emergency ram air valve provides a backup fresh air supply for unpressurized flight below 10,000 ft. The emergency ram air valve is powered by DC ESS BUS 3 and opens when the RAM AIR PBA is pressed. The emergency ram air valve connects the ram air inlet to the right pack discharge duct. If bleed air is available, the ram air can be mixed with hot air from the trim air system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-56
CS130_21_L2_RAS_General.fm
21 - Environmental Control 21-52 Ram Air System
DC ESS BUS 3
ICCP RAM AIR
OPEN
DMC 1 and DMC 2
RIGHT BULKHEAD CHECK VALVE Pack Discharge to Mix Manifold RIGHT PACK
M
EMERGENCY RAM AIR VALVE Ram Air Inlet
CS1_CS3_2150_023
LEGEND
Cold Air DMC Data Concentrator Unit Module Cabinet ICCP Integrated Cockpit Control Panel ARINC 429
Figure 24: Ram Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-57
CS130_21_L2_RAS_General.fm
21 - Environmental Control 21-52 Ram Air System
COMPONENT LOCATION The following components are located in the ram air system: •
Emergency ram air valve
EMERGENCY RAM AIR VALVE The emergency ram air valve is installed between the right pack discharge duct and the ram air inlet duct.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-58
CS130_21_L2_RAS_ComponentLocation.fm
21 - Environmental Control 21-52 Ram Air System
CS1_CS3_2120_017
Emergency Ram Air Valve
Figure 25: Emergency Ram Air Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-59
CS130_21_L2_RAS_ComponentLocation.fm
21 - Environmental Control 21-52 Ram Air System
CONTROLS AND INDICATIONS The RAM AIR PBA opens the emergency ram air valve. An OPEN legend on the RAM AIR PBA indicates the valve is open. The RAM AIR valve and flowbar on the AIR synoptic page illuminate green when the ram air system is selected ON.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-60
CS130_21_L2_RAS_ControlsIndications.fm
21 - Environmental Control 21-52 Ram Air System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
AIR COCKPIT
C
FWD
H
CABIN
C
AFT
H
FWD
MAN TEMP
CARGO
VENT
OFF
C
LO HEAT
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°C
10 °C
15 °C
H
15 °C RAM AIR
TRIM AIR
AFT
OFF
VENT
R PACK
L PACK
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
FAIL
PSI
XBLEED
45 PSI
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
45
OPEN
FAIL OFF R BLEED FAIL
APU BLEED
OFF
APU
OFF
AIR PANEL SYNOPTIC PAGE – AIR
CS1_CS3_2150_024
FAIL
Figure 26: Ram Air System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-61
CS130_21_L2_RAS_ControlsIndications.fm
21 - Environmental Control 21-52 Ram Air System
DETAILED DESCRIPTION The emergency ram air ventilation system supplies fresh airflow for unpressurized operation. The system supplies fresh air at altitudes up to 10,000 ft. The aircraft speed needs to be maintained above 0.5 Mach to ensure that enough dynamic air pressure is available at the ram air inlet. The emergency ram air valve connects the right hand pack ram air inlet diffuser to the right hand pack discharge duct. Air flows into the mix manifold for distribution throughout the aircraft. The valve has a manual override with a visual position indication. Microswitches provide position feedback to the data concentrator unit module cabinets (DMCs). The emergency ram air valve is powered from DC ESSENTIAL BUS 3 supplied through the hardwired RAM AIR guarded PBA on the AIR panel. An OPEN legend on the RAM AIR PBA, and a RAM AIR OPEN status message indicate the valve is open. When the RAM AIR PBA is selected, the outflow valve is commanded open to prevent excessive cabin pressure differential pressure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L3_RAS_DetailedDescription.fm
21 - Environmental Control 21-52 Ram Air System
ICCP Ram Air PBA
EMERGENCY RAM AIR VALVE Close
DC ESS BUS 3
3 Open Emer Ram Air Vlv Position
PIM
DMC 1 and DMC 2
Outflow Valve Open Command DMC ICCP PIM
CS1_CS3_2150_008
LEGEND Data Concentrator Unit Module Cabinet Integrated Cockpit Control Panel Panel Interface Module ARINC 429
Figure 27: Ram Air System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-63
CS130_21_L3_RAS_DetailedDescription.fm
21 - Environmental Control 21-52 Ram Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the ram air system:
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-64
CS130_21_L2_RAS_MonitoringTests.fm
21 - Environmental Control 21-52 Ram Air System
CAS MESSAGES Table 10: CAUTION Message MESSAGE RAM AIR FAIL
LOGIC RAM AIR selected OPEN/CLOSED and ERAV not detected full open/full closed.
Table 11: STATUS Message MESSAGE RAM AIR OPEN
LOGIC RAM AIR selected and confirmed OPEN.
Table 12: INFO Messages MESSAGE
LOGIC
21 AIR SYSTEM ERAV in unknown position. FAULT - ERAV INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_RAS_MonitoringTests.fm
21 - Environmental Control 21-60 Trim Air System
21-60 TRIM AIR SYSTEM GENERAL DESCRIPTION The trim air system is used for independent zone temperature control. The aircraft zones serviced by the trim air are the flight deck, forward cabin, aft cabin, and forward cargo compartment. Bleed air is supplied from each pack inlet line upstream of the flow control valves. The trim air shutoff valve (TASOV) on the left side, and the trim air pressure-regulating valve (TAPRV) on the right side supply bleed air to the trim air system. The trim system takes hot air upstream the flow control valve and ducts it to the distribution system downstream of the mix manifold. The trim air is injected in the individual zone distribution air supply ducting, and mixed with cold air from packs to provide the required temperature for the zone. Each trim air line is fitted with a trim air check valve (TACKV), installed in line at the pressure bulkhead. In case of a trim air line disconnect it isolates the pressurized cabin from the unpressurized wing-to-body faring (WTBF) bay of the aircraft to prevent cabin depressurization.
In normal operation, the system is on and controlled automatically by the IASCs. The TAPRV modulates to provide 12.8 psi pressure, to the TAVs. The TASOV remains closed during normal operation. If the TAPRV fails closed, the TASOV is driven open by IASC 1 and used to supply the trim air system. To improve heating performance on the ground during the flow control pull up mode, the TASOV is opened by IASC 1. If the trim air system is selected off using the TRIM AIR PBA on the AIR panel or commanded closed when the flow control valves (FCVs) close, the TASOV, TAPRV, and TAVs close. The flight deck TAV remains slightly open to prevent a pressure build-up in the trim air distribution lines.
The TASOV and trim air valves (TAVs) are powered through their respective integrated air system controller (IASC). IASC 1 powers: •
TASOV
•
Flight deck TAV
•
Aft cabin TAV
IASC 2 powers: •
Forward cabin TAV
•
Forward cargo TAV
The TAPRV is powered by DC ESS BUS 1, and reports its position to IACSC 2. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-66
CS130_21_L2_TAS_General.fm
21 - Environmental Control 21-60 Trim Air System
LEGEND
CDC DMC
Bleed Air Cold Air Recirculation Air Conditioned Air Control and Distribution Cabinet Data Concentrator Unit Module Cabinet
FCV IASC
Flow Control Valve Integrated Air System Controller
TAPRV TASOV TAV
C
H
C
H
OFF
VENT
CARGO LO HEAT
OFF
HI HEAT
ON
IASC 1
Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve ARINC 429
C
FWD
MAN TEMP
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
IASC 2
DMC 1 and DMC 2
DC ESS BUS 1 CDC 2
LEFT PACK
From Bleed Air System
FLIGHT DECK
L FCV FWD CABIN EMERGENCY RAM AIR VALVE
RIGHT PACK
From Bleed Air System R FCV
MIX MANIFOLD
M
M
TASOV
FLIGHT DECK TAV
AFT CABIN
FWD CARGO COMPARTMENT M
M
TAPRV
AFT CABIN TAV
M FWD CARGO TAV
CS1_CS3_2160_016
FWD CABIN TAV SOL
Figure 28: Trim Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-67
CS130_21_L2_TAS_General.fm
21 - Environmental Control 21-60 Trim Air System
COMPONENT LOCATION The following components are located in the WTBF: •
Trim air pressure-regulating valve (TAPRV)
•
Trim air shutoff valve (TASOV)
•
Trim air check valve (TACKV)
The following components are located in the environmental control system bay: •
Trim air valves (TAVs) (Refer to figure 30)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-68
CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
Trim Air Check Valves
B
B TRIM AIR PRESSURE-REGULATING VALVE
A TRIM AIR SHUTOFF VALVE
CS1_CS3_2160_007
A
Figure 29: Trim Air Shutoff Valves, Trim Air Check Valves, and Trim Air Pressure-Regulating Valve Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
TRIM AIR VALVES The trim air valves (TAVs) are located along the trim air distribution line in the environmental control system bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21-70
CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
Trim Air Distribution Line
Aft Cabin Trim Air Valve
Flight Deck Trim Air Valve
Fwd Cabin Trim Air Valve
TRIM AIR VALVE
CS1_CS3_2160_008
Cargo Trim Air Valve
Figure 30: Trim Air Valve Locations Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-71
CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
CONTROLS AND INDICATIONS The TRIM AIR PBA turns off the trim air system. An OFF legend on the TRIM AIR PBA indicates the trim air system is off. A TRIM AIR OFF message appears on the AIR synoptic page to indicate the system is off.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TAS_ControlsIndications.fm
21 - Environmental Control 21-60 Trim Air System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°C
10 °C
15 °C
15 °C
AIR COCKPIT
FWD
CABIN
AFT
R PACK
L PACK C
H
C
H
C
FWD
MAN TEMP OFF
VENT
CARGO LO HEAT
H
AFT
OFF
VENT
HI HEAT
OFF
45 RAM AIR TRIM AIR
RAM AIR
TRIM AIR
PACK FLOW
PSI
45
XBLEED
PSI
RECIRC AIR
OFF
APU
SYNOPTIC PAGE – AIR
CS1_CS3_2160_005
AIR PANEL
Figure 31: Trim Air System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TAS_ControlsIndications.fm
21 - Environmental Control 21-60 Trim Air System
DETAILED DESCRIPTION TRIM AIR CONTROL Each trim air valve (TAV) is commanded by the IASC according to the duct temperature sensor (DTS) measurement compared with a reference temperature, which is determined by the error between the selected temperature from the AIR panel in the flight deck and the temperature measured by the associated zone ventilated temperature sensor (VENTS). When the trim air system is selected OFF, the TAVs are also commanded closed by the IASCs. The flight deck TAV is commanded slightly open to avoid a pressure buildup in the trim air lines. The flight deck, forward cabin, aft cabin, and the FWD cargo TAV are automatically closed whenever the trim air system is closed, and either the TAPRV or the TASOV fails open.
•
Pack start up
•
Engine start + 30 seconds
•
Ditching
The cargo TAV is automatically closed when: •
The cargo heating system is selected OFF by the flight crew FWD CARGO switch set to OFF or VENT
•
The cargo SOVs are commanded closed due to forward cargo compartment smoke detection
•
TRIM AIR PBA OFF
•
Cargo DTS overheat
•
Total loss of cargo DTS
NOTE When the trim system is shutoff and all TAVs are closed, the flight deck TAV is commanded to open for 2.7 seconds, to prevent pressure build-up in the trim air distribution line. The TASOV and TAPRV are automatically closed whenever: •
DTS overheat is detected
•
Bleed leak is detected affecting the TAPRV and TASOV downstream ducting
•
Both flow control valves (FCVs) are commanded closed, and the emergency ram air valve (ERAV) is not selected open
•
TRIM AIR PBA is selected OFF
•
The TAPRV or the TASOV are commanded closed when its associated FCV is closed, and ERAV is not selected open.
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21 - Environmental Control 21-60 Trim Air System
LEGEND Bleed Air Cold Air Recirculation Air Conditioned Air Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Duct Temperature Sensor Integrated Air System Controller Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve Ventilated Temperature Sensor ARINC 429 Discrete
Recirculation Air
From Left Pack
MIXING MANIFOLD
From Right Pack
DTS
FLIGHT DECK VENTS
CDC DMC DTS IASC TAPRV TASOV TAV VENTS
Recirculation Air IASC 1
Channel B COCKPIT
From Bleed Air System From Bleed Air System
M
FLIGHT DECK TRIM AIR VALVE
M SOL
TAPRV
To Other Trim Air Valves
+ -
Flight Deck Duct Reference Channel A
NOTE
+
C
CABIN
H
C
CARGO
MAN TEMP
HEAT
-
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
AIR PANEL
Flight deck shown, FWD cabin, aft cabin, and FWD cargo similar.
AFT
OPEN
CS1_CS3_2150_017
TASOV
Figure 32: Trim Air Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_TAS_DetailedDescription.fm
21 - Environmental Control 21-60 Trim Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the trim air system:
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21 - Environmental Control 21-60 Trim Air System
CAS MESSAGES Table 16: INFO Messages
Table 13: CAUTION Messages MESSAGE
LOGIC
MESSAGE
LOGIC
TRIM AIR FAIL
Trim air failure or loss of trim air leak detection or loss of both IASC 1channels.
FWD CARGO HEAT FAIL
LO HEAT AND HI HEAT mode not available.
21 AIR SYSTEM TAPRV failed closed. FAULT - TRIM AIR PRV FAIL CLSD
FWD CARGO LO TEMP
Low temperature <4°C in fwd cargo when FWD CARGO selected to LO/HI HEAT
21 AIR SYSTEM TAPRV failed open. FAULT - TRIM AIR PRV FAIL OPEN 21 AIR SYSTEM TASOV failed closed in full closed position. FAULT - TRIM AIR SOV FAIL CLSD
Table 14: ADVISORY Message MESSAGE AIR SYSTEM FAULT
LOGIC Fault not requiring immediate attention. (Total loss of VENTS, IASC single-channel failure, loss of cockpit, FWD, AFT cabin DTS, TAV stuck, TAPRV/TASOV failed closed, single channel CAR DTS, or CATS failed).
Table 15: STATUS Messages MESSAGE TRIM AIR OFF
LOGIC TRIM AIR selected OFF and both TAPRV and TASOV confirmed closed.
FWD CARGO AIR FWD cargo heating selection selected and confirmed OFF OFF. AFT CARGO AIR OFF
AFT cargo heating selection selected and confirmed OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
21 AIR SYSTEM TASOV failed not closed or unknown position. FAULT - TRIM AIR SOV FAIL OPEN 21 AIR SYSTEM FAULT - FWD CARGO SOV INOP
Upstream or downstream fwd cargo SOV failed in position or in unknown position.
21 AIR SYSTEM FAULT - AFT CARGO SOV INOP
Upstream or downstream aft cargo SOV failed.
21 AIR SYSTEM FAULT - FWD CAR TAV FAIL CLSD
CAR TAV failed closed.
21 AIR SYSTEM Loss of CKPT or FWD CAB or AFT CAB TAV or any TAV FAULT - TAV INOP unknown position
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21 - Environmental Control 21-60 Temperature Control System
21-60 TEMPERATURE CONTROL SYSTEM GENERAL DESCRIPTION The temperature control system controls the flight deck, forward and aft cabin, and the forward cargo compartment. Individual automatic temperature controls for the flight deck and forward and aft cabins are available on the AIR panel. The forward and aft cabin temperature is also controlled from the cabin management system (CMS) crew terminal. The forward cargo heat is selectable from the AIR panel. The system temperature control for the flight deck, forward and aft cabins, and forward cargo compartment is automatically controlled via the integrated air system controllers (IASCs) in order to meet the flight deck temperature selections. The IASCs receive temperature inputs from: •
Ventilated temperature sensor for actual zone temperature
•
Duct temperature sensors for duct temperature
•
Mix manifold temperature sensors for pack temperature control
•
Cargo temperature sensor for forward cargo actual temperature
•
Cargo duct temperature sensor for cargo duct temperature
The forward and aft cabin temperature can be adjusted at the cabin management system crew terminal. The crew terminal provides the capability to adjust the temperature by ± 3°C (5.4°F) from the temperature selection made on the AIR panel FWD, or AFT CABIN temperature selectors. The forward cargo LO HEAT provides a temperature range of 15°C (59°F) to 20°C (68°F), and HI HEAT provides a temperature range of 20°C (68°F) to 25°C (77°F).
The packs are controlled to provide the lowest temperature output required by the flight deck, forward, or aft cabin zones. The IASCs control the trim air valve (TAV) to add hot air to each of the distribution ducts that require additional heating. In automatic mode, the required zone temperature for the flight deck, forward, and aft cabin can be set from 18°C (64°F) to 30°C (86°F) using the AIR panel temperature selectors. In manual mode, the flight crew adjust the zone temperatures by adjusting the duct temperature from full cold 5°C (41°F) to full hot 65°C (149°F) using the AIR panel temperature selectors.
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CS130_21_L2_TCS_General.fm
21 - Environmental Control 21-60 Temperature Control System
LEGEND Hot Air Cold Air Recirculation Air Conditioned Air Cargo Duct Temperature Sensor Cargo Temperature Sensor Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Duct Temperature Sensor Integrated Air System Controller Mix Manifold Temperature Sensor Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve Ventilated Temperature Sensor ARINC 429
C
FWD
H
C
H
FWD
MAN TEMP OFF
CABIN
VENT
AFT
C
CARGO LO HEAT
CABIN MANAGEMENT SYSTEM CREW TERMINAL
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
OPEN
DMC 1 and DMC 2
IASC1
IASC2
CDC 2
DTS
CADTS CATS CDC DMC DTS IASC MIXTS TAPRV TASOV TAV VENTS
AIR COCKPIT
From Left Pack
UNDERFLOOR CATS
DTS
MIX MANIFOLD
MIXTS 2
From Bleed Air System
TASOV M
TAPRV
SOL
VENTS
FWD CARGO COMPARTMENT
M AFT CABIN TAV FWD CARGO TAV M
FWD CABIN
AFT CABIN
M F/D TAV M FWD CABIN TAV
VENTS
CADTS
CS1_CS3_2160_021
From Right Pack
From Bleed Air System
FLIGHT DECK DTS
MIXTS 1
VENTS
Figure 33: Temperature Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-79
CS130_21_L2_TCS_General.fm
21 - Environmental Control 21-60 Temperature Control System
COMPONENT LOCATION The following are major components of the temperature control system: •
Mix manifold temperature sensors (MIXTS)
•
Duct temperature sensor (DTS) (Refer to figure 35)
•
Ventilated temperature sensors (VENTS) (Refer to figure 36)
•
Cargo duct temperature sensor (CADTS) (Refer to figure 37)
•
Cargo temperature sensor (CATS) (Refer to figure 37)
MIX MANIFOLD TEMPERATURE SENSORS The mix manifold temperature sensors are located in the pack discharge lines near the inlet to the mix manifold.
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21 - Environmental Control 21-60 Temperature Control System
Mix Manifold Temperature Sensors
CS1_CS3_2120_029
TEMPERATURE SENSOR
Figure 34: Mix Manifold Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-81
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
DUCT TEMPERATURE SENSORS The duct temperature sensors (DTS) for the forward cabin, aft cabin, and flight deck are located in the environmental control system bay.
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CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
Fwd Cabin Duct Temperature Sensor
Aft Cabin Duct Temperature Sensor TEMPERATURE SENSOR
CS1_CS3_2160_011
Flight Deck Duct Temperature Sensor
Figure 35: Duct Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
VENTILATED TEMPERATURE SENSORS Flight Deck Ventilated Temperature Sensor The flight deck ventilated temperature sensor is installed on the pilots bulkhead. Forward Cabin Ventilated Temperature Sensor The forward cabin ventilated temperature sensor is located on the right side of the forward cabin. The sensor is mounted in the passenger service unit located between frame 34 and frame 35. Aft Cabin Ventilated Temperature Sensor The aft cabin ventilated temperature sensor is located on the left side of the aft cabin. The sensor is mounted in the passenger service unit located between frame 62 and frame 64.
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21 - Environmental Control 21-60 Temperature Control System
C
A
B
A
VENTILATED TEMPERATURE SENSOR
Ventilated Temperature Sensor
B AFT CABIN
C FWD CABIN
CS1_CS3_2160_010
Ventilated Temperature Sensor
Figure 36: Flight Deck and Cabin Ventilated Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-85
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
CARGO DUCT TEMPERATURE SENSOR The cargo duct temperature sensor (CADTS) is located in the forward cargo ventilation inlet duct in the environmental control system (ECS) bay. CARGO TEMPERATURE SENSOR The cargo temperature sensor (CATS) is located in the forward cargo ventilation exhaust duct in the environmental control system bay.
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CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
Fwd Cargo Compartment
TEMPERATURE SENSOR Cargo Temperature Sensor
CS1_CS3_2160_009
Cargo Duct Temperature Sensor
Figure 37: Cargo Duct Temperature Sensor and Cargo Temperature Sensor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-87
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
COMPONENT INFORMATION VENTILATED TEMPERATURE SENSOR The ventilated temperature sensor (VENTS) supply temperature information to the IASC. The vents consist of two temperature sensing elements installed on a circuit board. The circuit board is housed in a plastic box along with a small fan. The fan draws air across the temperature sensor. A filter is installed on the cover to ensure the temperature sensors stay clean.
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CS130_21_L2_TCS_ComponentInformation.fm
21 - Environmental Control 21-60 Temperature Control System
CS1_CS3_2160_022
Filter
Figure 38: Ventilated Temperature Sensor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-89
CS130_21_L2_TCS_ComponentInformation.fm
21 - Environmental Control 21-60 Temperature Control System
CONTROLS AND INDICATIONS AIR PANEL The COCKPIT, FWD and AFT CABIN temperature control selectors are used to set the desired temperature in automatic and manual modes. A MAN TEMP PBA is to be used to select manual control of the temperature control system. An ON legend indicates the system is in manual control. The FWD CARGO rotary switch is used to select LO HEAT or HI HEAT in the forward cargo compartment.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
AIR COCKPIT
C
FWD
H
CABIN
C
H
C
FWD
MAN TEMP
CARGO
VENT
OFF
AFT
LO HEAT
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR PACK FLOW
RECIRC AIR
L PACK
R PACK AIR PANEL
CS1_CS3_2160_020
TRIM AIR
Figure 39: Temperature System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-91
CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
AIR SYNOPTIC PAGE The AIR synoptic page indicates the actual temperature for the flight deck, forward, and aft cabins on top. The selected temperature from the AIR panel is shown below the actual temperature. If the selected temperature has been modified by the cabin management system (CMS) crew terminal temperature page selection, that information is displayed beside the selected temperature. The duct temperature is indicated at the outlet of the mix manifold representation on the synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
STATUSS
AIR IR
DOOR OR
ELECC
FLT CTR CTRLL
FUEL FUE EL
HYD YD
AVIONIC
INFO
CB
AIR SYNOPTIC PAGE INDICATIONS
2
3
CKPT 00 °C 24 °C
24°CC
1
CARGO
LO 15 °C °
FWD 26 °C4 24 °°C -+ 2
AFT 26 °C 24 °C -+ 2
24 °CC
Fwd Cargo Selection Indicator: • OFF – no heating, no ventilation • VENT – cargo ventilation only • LO HEAT – vent and low heating level (15-20°C) • HI HEAT – vent and warmer heating level (20-25°C)
1
CARGO LO
2
AFT 26°C
Actual Cabin Temperature: • Dashed line "ííí" represents a temperature sensor failure or operating in manual mode
3
24°C
Selected Temperature: • Automatic mode – desired zone temperature • Manual mode – desired duct temperature
4
+2
Cabin Management System Input (+/-3°C)
5
24°C
Duct Temperature
24 °C 5
RAM AIR
R PACK
L PACK
CS1_CS3_2160_018
SYNOPTIC PAGE – AIR
Figure 40: Temperature System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
CABIN MANAGEMENT SYSTEM TEMPERATURE CONTROLS The current temperature levels, and the temperature adjustment for each cabin zone are displayed on the temperature screen. The allowable range is displayed for each cabin zone. The current temperature in each cabin is displayed on a layout of passenger area (LOPA) screen. If the temperature adjustment is not available, the temperature controls are disabled and an error message is displayed on the screen.
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21 - Environmental Control
CS1_CS3_2160_019
21-60 Temperature Control System
Figure 41: Cabin Management System Temperature Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-95
CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
DETAILED DESCRIPTION AUTOMATIC TEMPERATURE CONTROL The automatic temperature control system controls the air conditioning packs, ram air regulating valve (RARV), recirculation system, and the trim air system to provide the selected temperature for the flight deck, the forward, and aft cabins, and the forward cargo compartment. For zone temperature control, the integrated air system controllers (IASCs) use the following sensors: •
The ventilated temperature sensors (VENTS), located in the flight deck, forward cabin and aft cabin zones sense the actual zone temperature
•
The duct temperature sensors (DTS), located in the main zone distribution lines downstream of the TAV, sense the temperature of the air supplied to each zone
•
The mix temperature sensors (MIXTS) sense the temperature of the pack discharge and recirculation mixed air for the pack temperature control valve (TCV)
•
The pack discharge temperature sensors (PDTS), located on pack discharge ducts, sense each pack discharge temperature. The PDTS is mainly used as a monitoring means, such as pack discharge overheat protection or as a backup to the MIXTS, the PTS or the PDPS for redundant icing protection.
The IASC modulates the TCV and RARV position to achieve the reference mix manifold temperature. The mix manifold reference temperature is set equal to the lowest reference temperature demand computed for each of the three independent zones. The IASC controls each TAV in a closed loop against the duct temperature sensor measurement. The DTS reference temperature is computed from the difference between the ventilated temperature sensor measurement and the zone temperature selections made on the AIR panel.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Each TAV is commanded by its IASC according to the duct temperature sensor (DTS) measurement compared with a reference temperature determined by the error between the selected temperature from the AIR panel in the flight deck, and the temperature measured by the associated zone ventilated temperature sensor (VENTS). •
If the forward or aft cabin VENTS fails, the remaining VENTS is used for temperature control in both cabins
•
If the flight deck VENTS or both the forward and aft cabin VENTS fails, the affected zones are controlled using the supply duct temperature as a reference similar to the manual temperature control
•
If a temperature selector on the AIR panel fails, the system assumes a default value of 24°C (75°F) for that zone
•
If the forward or aft DTS fails, the associated TAV is closed and the remaining zone provides temperature control for the entire cabin
•
If the flight deck DTS fails, the TRIM AIR PBA must be selected OFF
If the flight deck, forward or aft cabin DTS measurement exceeds 85°C (185°F) for over 30 seconds, the trim air system is closed automatically, and a TRIM AIR FAIL caution message is posted and latched until the TRIM AIR PBA is selected OFF, and the overheat condition disappears. MANUAL TEMPERATURE CONTROL In manual mode, the flight crew can manually adjust the zone temperatures from the AIR panel by selecting the MAN TEMP PBA to ON. The duct temperature sensors (DTS) are used for the reference temperature. The temperature in the duct can be controlled from 5°C (41°F) to 65°C (149°F). When in manual temperature control, the selected temperature reading on the synoptic page becomes the DTS target, and the actual temperature displayed on the synoptic page becomes dashed.
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21 - Environmental Control 21-60 Temperature Control System
Ram Air Inlet
AIR COCKPIT
C
FWD
H
C
CABIN
H
FWD
MAN TEMP OFF
VENT
PTS
CDTS
H
AFT
OFF
VENT
HI HEAT
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
-
Hot Air Warm Air Cold Air Water Duct Temperature Sensor Integrated Air System Controller Minimum Mix Manifold Temperature Sensor Pack Discharge Temperature Sensor Pack Temperature Sensor Ram Air Regulating Valve Trim Air Valve Temperature Control Valve IASC Ventilated Temperature Sensor
+
VENTS
F/D TAV M
Temperature Control Valve
LEGEND
FLIGHT DECK
+
-
-
Flight Deck Duct Reference
Trim Reference
MIXTS Reference
+
Min Temp
+
FWD Cabin Duct Reference
-
Trim Air Available
+
Aft Cabin Duct Reference
-
FWD Cabin Vents
Aft Cabin Vents
CS1_CS3_2161_001
Ram Air Regulating Valve
MIX MANIFOLD
To PDPS
M
M
MIXTS
PDTS
Condenser
DTS
TURB
COMP
Reheater
Air Cycle Machine
DTS IASC MIN MIXTS PDTS PTS RARV TAV TCV VENTS
C
CARGO LO HEAT
ON
From Bleed Air System
AFT
Figure 42: Flight Deck, Forward Cabin, and Aft Cabin Temperature Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_TCS_DetailedDescription.fm
21 - Environmental Control 21-60 Temperature Control System
FWD CARGO TEMPERATURE CONTROL The TAV mixes hot air to the flow distributed from the recirculation system to the forward cargo in order to provide the required temperature. The temperature is controlled in a closed loop by measurement feedback from the cargo temperature sensor (CATS) to IASC 2. The cargo duct temperature sensor (CADTS) provides temperature measurement to IASC 2 for overheat monitoring and protection. A zone target temperature is set, based on LO or HI heat selection and the duct temperature reference, based on the error between this target and the measured temperature. If the cargo duct temperature exceeds 70°C (158°F) for 30 seconds, the trim air system is closed automatically. A TRIM AIR FAIL caution message is displayed until the TRIM AIR PBA is selected OFF, and the overheat condition disappears. If the CADTS or CATS fails, the FWD CARGO HEAT is selected OFF. When the FWD cargo selector is set to OFF or VENT, and the CARGO TAV full closed switch is not received, a TRIM AIR FAIL caution message is posted and latched until the TRIM AIR PBA is selected OFF.
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21 - Environmental Control 21-60 Temperature Control System
CATS DOWNSTREAM CARGO SOV FWD CARGO COMPARTMENT IASC2 UPSTREAM CARGO SOV
M From Trim Air System
+
FWD Cargo Duct Reference
CADTS
Recirculation Air
FWD + -
OFF
VENT
CA LO HEAT HI HEAT
AIR PANEL
FWD CARGO TAV
LEGEND
CS1_CS3_2161_002
CADTS CATS IASC SOV TAV
Hot Air Warm Air Cold Air Water Cargo Duct Temperature Sensor Cargo Temperature Sensor Integrated Air System Controller Shutoff valve Trim Air Valve
Figure 43: Cargo Temperature Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-60 Temperature Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the temperature control system:
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21 - Environmental Control 21-60 Temperature Control System
CAS MESSAGES Table 20: INFO Messages
Table 17: CAUTION Messages MESSAGE
LOGIC
MESSAGE
FWD HEAT CARGO FAIL
Total loss of CADTS or CATS.
FWD CARGO LO TEMP
In the event the FWD cargo is selected to LO or HI HEAT, and the CATS reads below 10°C for at least 2 minutes during cruise.
Table 18: ADVISORY Messages MESSAGE AIR SYSTEM FAULT
PACK FAULT
LOGIC Fault not requiring immediate attention. (Total loss of VENTS, IASC single channel failure, loss of cockpit, FWD, AFT cabin DTS, TAV stuck, TAPRV/TASOV failed closed, single-channel CADTS, or CATS failed). Loss of redundant or non-critical function for the air conditioning system.
LOGIC
21 AIR SYSTEM FAULT - ZONE TEMP SNSR INOP
Cockpit VENTS total loss (both channels) OR
21 AIR SYSTEM FAULT - DUCT TEMP SNSR INOP
Total loss or drift of FWD CABIN DTS or AFT CABIN DTS or redundant loss or drift of CKPT DTS.
21 PACK FAULT MIX MANF TEMP SNSR TOTAL LOSS
MMTS from both packs are lost.
21 PACK FAULT MIX MANF TEMP SNSR REDUND LOSS
Up to three MMTS sensing elements redundant loss.
Aft Cabin VENTS total loss (both channels) OR FWD Cabin VENTS total loss (both channels).
Table 19: STATUS Messages MESSAGE
LOGIC
FWD CARGO AIR FWD cargo heating selection selected and confirmed OFF OFF. AFT CARGO AIR OFF
AFT cargo heating selection selected and confirmed OFF.
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21 - Environmental Control 21-22 Recirculation
21-22 RECIRCULATION GENERAL DESCRIPTION Cabin air is recycled into the mixing manifold by mean of a single recirculation fan (RFAN), which draws air from the cabin underfloor across two air filters. The RFAN supplies the air into each pack discharge line for initial mixing with the pack discharge air, downstream of the bulkhead check valve. Integrated air system controller (IASC) 1 controls the RFAN. If IASC 1 fails, IASC 2 controls the recirculation fan. The recirculation fan has a motor controller that the IASC signals to adjust the fan speed based mainly on cabin altitude. The fan is powered by AC BUS 2, and the controller is powered by DC BUS 2. The recirculation fan operation is fully automatic and runs when power is available, no faults exist, and there is no smoke in the equipment bays. The recirculation fan can be selected OFF using the RECIRC AIR PBA on the AIR panel. The RFAN is normally selected OFF when the fire detection and extinguishing (FIDEX) control unit reports smoke in the equipment bays. The recirculation fan is commanded so that the recirculation flow ratio remains between 25% and 40% of the total air flow for normal operation, and between 35% and 50% of the total airflow for single pack operation. The recirculation flow also ensures the mixing manifolds temperature remains above freezing.
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21 - Environmental Control 21-22 Recirculation
From Left Hand Pack
From Right Hand Pack
Low Pressure Ground Connection
Emergency Ram Air Valve
Pressurized Cabin Underfloor
Pressurized Cabin Underfloor
Check Valves
Cargo Ventilation MIXING MANIFOLD
RECIRCULATION FILTER
RECIRCULATION FILTER
RECIRC AIR FIDEX CONTROL UNIT
DMC 1 and DMC 2
RECIRCULATION FAN
OFF
AC BUS 2 DC BUS 2
DMC IASC
Cold Air Recirculated Air Data Concentrator Unit Module Cabinet Integrated Air System Controller ARINC 429 CAN BUS
IASC 1 and IASC 2
CS1_CS3_2120_019
LEGEND
Figure 44: Recirculation System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
COMPONENT LOCATION The following components are located in the environmental control system bay, aft of the forward cargo compartment: •
Recirculation fan filters
•
Recirculation fan and motor controller
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21 - Environmental Control 21-22 Recirculation
CS1_CS3_2120_012
Recirculation Fan Filters
Recirculation Fan and Motor Controller
Figure 45: Recirculation System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
COMPONENT INFORMATION RECIRCULATION FILTERS The disposable recirculation filter ensures cleaned air is recycled to the cabin. It is composed of an HEPA matrix, and an active charcoal core containing chemical absorbers used to control odors. The recirculation filter is supported by a structure at the bottom and at the top. A V-band clamp holds the filter at the bottom, and a plunger centers the filter at the top.
WARNING DO NOT TOUCH THE FILTER WITH BARE HANDS. USE GLOVES AND PERSONAL PROTECTIVE EQUIPMENT TO HANDLE THE FILTER. THE FILTER SHOULD BE CONSIDERED A BIOHAZARD. THE FILTER CAN CONTAIN GERMS AND VIRUSES THAT CAN CAUSE SICKNESS.
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21 - Environmental Control
CS1_CS3_2120_026
21-22 Recirculation
Figure 46: Recirculation Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
CONTROLS AND INDICATIONS The RECIRC AIR PBA turns off the recirculation fan. An OFF legend on the RECIRC AIR PBA indicates the system is shutdown. The RECIRC AIR PBA is used when there is a recirculation fan failure or smoke is detected in the avionics cooling air extraction system. The PBA is selected to match the system configuration. A RECIRC OFF message appears on the AIR synoptic page to indicate the system is shutdown.
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21 - Environmental Control 21-22 Recirculation
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°CC
10 °°CC
15 °CC
15 °CC
AIR COCKPIT
FWD
CABIN
AFT
RAM AIR
TRIM AIR R PACK
L PACK C
H
C
H
C
FWD
MAN TEMP OFF
VENT
CARGO LO HEAT
RECIRC OFF
H
AFT
OFF
VENT
HI HEAT
45 RAM AIR TRIM AIR
PACK FLOW
PSI
XBLEED
45 PSI
RECIRC AIR OFF
APU
SYNOPTIC PAGE – AIR
CS1_CS3_2120_025
AIR PANEL
Figure 47: Recirculation System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
DETAILED DESCRIPTION The RFAN 115 VAC power is supplied from AC BUS 2 through a relay in electrical power center (EPC 2). The relay is controlled by control and distribution cabinet 2 (CDC 2). CDC 2 monitors the relay position using the discrete feedback from a relay contact. The position information is sent to the IASCs through the data concentrator unit module cabinet (DMC). The recirculation system operation is fully automatic. The RECIRC AIR PBA is used only in case of recirculation fan failure or in case of smoke detected in the avionics extraction system, and confirms shutdown of the recirculation system. The PBA commands the CDC 2 to isolate the 115 VAC power from the RFAN, bypassing IASC commands. Each IASC communicates with the RFAN controller through a CAN BUS to automatically adjust the fan speed. The IASC controls the RFAN speed according to a schedule as a function of system configuration •
Number of packs operating
•
Pull-up or pull-down sequence
•
Cabin altitude
•
Weight-on-wheels (WOW) signal
•
Cargo ventilation status
•
Aircraft version (CS100 or CS300)
If the FWD and AFT CARGO switches are selected OFF, with the aircraft operating on ground at a field elevation greater than 12,000 ft, the recirculation fan is automatically turned off by the IACSs to prevent excessive distribution noise. The normal maximum fan speed is 15,800 rpm. If the cabin altitude is out of its normal range, the RFAN speed increases to 16,000 rpm. The RFAN provides protection against internal overtemperature, overcurrent, and overvoltage. In case of a major internal failure, the RFAN is automatically turned off, and latched until the RFAN power is recycled. Once the RFAN fault is detected by the IASCs either from the controller or from the solid-state power controller (SSPC) feedback, an RECIRC AIR FAIL caution message is displayed.
NOTE 1. On the CS100, the recirculation fan is shutoff during the pull-up sequence. 2. On the CS300, the recirculation fan runs at maximum speed during the pull-up sequence.
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21 - Environmental Control 21-22 Recirculation
Recirc Fan AC BUS 2 20A RECIRCULATION FAN FIDEX CONTROL UNIT
RECIRC AIR
EPC 2 DMC 1 AND DMC 2
OFF
CDC 2-9-13 RECIRC FAN RLY CMD DC SSPC 3A BUS 2 Logic: Advanced
Recirc Fan Discrete Air/Gnd Recirc Fan Fault
CDC DMC EPC FIDEX IASC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electrical Power Center Fire Detection and Extinguishing Integrated Air Systems Controller AFDX ARINC 429 CAN BUS
Logic: Always On CDC 2 IASC 1 and IASC 2
CS1_CS3_2120_021
LEGEND
CDC 2-9-14 RECIRC FAN CTRL DC SSPC 3A BUS 2
Figure 48: Recirculation Fan Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
MONITORING AND TEST The following page provides the crew alerting system (CAS) and INFO messages associated with the recirculation system:
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21 - Environmental Control 21-22 Recirculation
CAS MESSAGES Table 21: CAUTION Message MESSAGE
LOGIC
RECIRC AIR FAIL RFAN internal failure detected. RFAN digital bus communication failure detected. Loss of both IASC B channels with RFAN powered and IASC safety channels remain functional.
Table 22: STATUS Message MESSAGE
LOGIC
RECIRC AIR OFF Recirculation air selected OFF.
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21 - Environmental Control 21-22 Distribution System
21-22 DISTRIBUTION SYSTEM GENERAL DESCRIPTION The low-pressure distribution system consists of a network of rigid and flexible composite ducts, which collects air from the air supply sources, mixes fresh conditioned air with recirculation air in a mixing manifold unit. Air conditioning is normally supplied from two air conditioning packs through low-pressure distribution lines. Air conditioning can also be supplied from a low-pressure ground cart through the low-pressure ground connection. An emergency ram air valve connects the ram air inlet to the low-pressure distribution system just upstream the bulkhead check valves in flight. This provides ventilation to the cabin if both air conditioning packs fail. Two bulkhead check valves isolate the pressurized cabin from the bleed air and air conditioning packs installed in the unpressurized wing-to-body fairing (WTBF). In the event of a distribution line rupture outside of the pressurized compartment, the bulkhead check valves close to minimize pressure loss in the aircraft. Air from the air conditioning packs is mixed with recirculation air just before entering the mix manifold. The mix manifold has three main outlets which split the flow for distribution. The air is distributed to three zones; flight deck, forward cabin, and aft cabin through dedicated distribution lines.
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21 - Environmental Control 21-22 Distribution System
Low-Pressure Ground Connection
From Left Pack
Emergency Ram Air Valve
From Right Pack
Ram Air
Check Valves
To Cargo Ventilation MIX MANIFOLD
To Aft Cabin
To Flight Deck Air To Forward Cabin
From Recirculation Fan
CS1_CS3_2120_024
From Recirculation Fan
To Gasper Air
LEGEND Cold Air Recirculated Air
Figure 49: Distribution System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
ZONE DISTRIBUTION
Lavatory and Galley Areas
Both forward and aft cabin zones and the flight deck are supplied independently from the mixing manifold, allowing for three independent zone temperature controls in the aircraft.
The air distribution is provided by air outlets located in the galley and lavatory areas. The ventilation flow for the galleys and lavatories is provided at a lower rate than the extraction flow to prevent odors from the galleys and lavatories flowing toward the passenger seating area.
The air is ducted by risers into overhead stowage bins on both sides of the cabin, and then discharged into a plenum that distributes air at each cabin row from the top, the bottom, and the side of the bins. On each side of the cabin, an underfloor distribution plenum connects the risers for the flight deck, the forward cabin, and the aft cabin with the mixing manifold. The end of each plenum connects with the galley and lavatory areas located at both ends of the cabin. Conditioned air is distributed via a dedicated gasper air line. The line runs from the mix manifold to risers on the right and left sides of the aircraft. The risers connect to the overhead stowage bins for distribution to the gasper air outlets on the passenger service units.
Flight attendant gasper air outlets in the galley and entryway areas are connected to the distribution lines. There are provisions for three gasper outlets for the forward zone, and four gasper outlets for the aft zone. Flight Deck The flight deck air is distributed by a dedicated duct line that runs along the left side underfloor of the cabin and supplies the flight deck. For the CS100, the flight deck/cabin flow split is 14%/86% respectively.
CS100 Cabin Distribution There are 10 risers on each side for the forward cabin zone. The farthest forward riser on each side is dedicated to the forward galley area. The aft riser on each side provides extra flow for the overwing emergency exit (OWEE) doors area ventilation. There are 12 risers on each side for the aft cabin zone. The farthest aft riser on each side is dedicated to the aft galley area. CS100 Gasper Air The CS100 has five gasper risers. The three on the right and two on the left connect the mix manifold gasper air outlet to the overhead stowage bins.
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21 - Environmental Control 21-22 Distribution System
Risers
Flight Deck Distribution Fwd Cabin Distribution Aft Cabin Distribution Gasper Air CS100
CS1_CS3_2120_015
LEGEND
Figure 50: Zone Distribution CS100 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
CS300 Cabin Distribution There are 12 risers on each side for the farthest forward cabin zone. The forward riser on each side is dedicated to the forward galley area. The most aft riser on each side provides extra flow for the overwing emergency exit (OWEE) doors area ventilation. There are 12 risers on each side for the aft cabin zone. The farthest aft riser on each side is dedicated to the aft galley area. For the CS300, the flight deck/cabin flow split is set at 12.5%/87.5% respectively. CS300 Gasper Air The CS300 has seven gasper risers. The four on the right and three on the left connect the mix manifold gasper air outlet to the overhead stowage bins.
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21 - Environmental Control 21-22 Distribution System
Risers
Flight Deck Distribution Fwd Cabin Distribution Aft Cabin Distribution Gasper Air CS300
CS1_CS3_2120_027
LEGEND
Figure 51: Zone Distribution CS300 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
OVERHEAD BIN DISTRIBUTION The forward and aft cabin air is distributed from the top of the cabin through airflow channels integral to the passenger baggage overhead bins. The low-pressure distribution network connects to the bins by means of duct risers located of both side of the inner fuselage.
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21 - Environmental Control 21-22 Distribution System
50%
GASPER AIR FLOW
From riser 25%
25%
CABIN AIR DISTRIBUTION
CS1_CS3_2120_028
From riser
Figure 52: Overhead Bin Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
FLIGHT DECK DISTRIBUTION From the mix manifold, air is routed to the flight deck via underfloor low-pressure ducting, passing into the flight deck within a single duct to provide ventilation for the crew via gaspers as well as general ventilation via diffusers in the side panels. Each gasper outlet is adjustable for flow and direction. The ventilation air is also passed into the flight deck through the glareshield. The air supplied to the flight deck is ducted up through the floor at the center console, and routed to provide additional ventilation behind the control panels. In the flight deck the following air outlets are provided: •
Large upward facing outlet in each side console
•
Gasper outlets above the pilot, copilot, and observer head ventilation
•
Gasper outlet on each side of the forward instrument panel
•
Gasper outlet on each side of the floor pan at foot level
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21 - Environmental Control 21-22 Distribution System
Overhead Gasper Outlets
Side Console Ventilation
Side Console Ventilation Fwd Glareshield Outlets
Flight Deck Distribution Duct
Knee Gasper Outlet
CS1_CS3_2120_032
Side Gasper Outlet
Figure 53: Flight Deck Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
COMPONENT LOCATION The following distribution system components are located in the environmental control system compartment, aft of the forward cargo compartment. •
Mix manifold
•
Bulkhead check valve
•
Low-pressure ground connection
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21 - Environmental Control 21-22 Distribution System
Mix Manifold
Bulkhead Check Valve
CS1_CS3_2120_030
Low-Pressure Ground Connection
Bulkhead Check Valve
Figure 54: Distribution System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
COMPONENT INFORMATION LOW-PRESSURE GROUND CONNECTION The low-pressure ground connection allows the connection of an external low-pressure ground cart to the aircraft during ground operations. The low-pressure ground cart supplies the aircraft cabin with conditioned air using the aircraft low-pressure distribution system. The low-pressure ground connection consists of an adapter, flapper check valve, and a cover. In flight, when the cabin is pressurized, the flapper check valve and adapter are designed to withstand differential pressure of 8.9 psid, which is slightly higher than the differential pressure seen by the pressurized vessel. A cover provides protection from water entering the main distribution system in the event of an aircraft ditching. The cover is held in place by a pin that remains attached to the aircraft by a lanyard.
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21 - Environmental Control 21-22 Distribution System
Check Valve
Pin Cover
CS1_CS3_2120_010
Adapter
Figure 55: Low-Pressure Ground Connection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
MIX MANIFOLD The mix manifold acts as a mixing chamber achieving, the required air temperature for distribution. Two air-conditioning packs feed fresh air into the mix manifold. This fresh air mixes with recirculated air collected from the underfloor area near the outflow valve. Some pre-mixing occurs prior to the mix manifold from the junction of the recirculation ducts and the pack ducts. The mix manifold has four main outlets, which split the flow for distribution to the forward cabin, aft cabin, flight deck, and gasper air.
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21 - Environmental Control 21-22 Distribution System
Mix Manifold Left Pack Air
Right Pack Air
Recirculation Air
Aft Cabin Distribution
Gasper Air
Flight Deck Distribution
CS1_CS3_2120_016
Fwd Cabin Distribution
Figure 56: Mix Manifold Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
21-23 VENTILATION SYSTEM GENERAL DESCRIPTION Air tapped off the recirculation air ventilates the cargo compartments through a cargo shutoff valve (CSOV). Each cargo compartment has two CSOVs. One is installed in the inlet duct, and the other is installed at the opposite end of the compartment in the exhaust duct. The CSOV is an electrically-actuated butterfly valve. Microswitches report the open and closed positions. The valve remains in its last position if no power is supplied. The airflow maintains the temperature above 4°C in the forward cargo, and 2°C in the aft cargo. The air is exhausted into the underfloor area surrounding the cargo compartments. The forward cargo compartment CSOV exhausts air close to the outflow valve to avoid recirculating odors when transporting live cargo. The valves are open for normal operation. When the recirculation fan is selected OFF, the integrated air system controllers (IASCs) close the CSOVs. If the forward cargo heat is selected, the forward CSOVs remain open, and air from the air conditioning units mix with cargo trim air. The CSOVs isolate the cargo compartment when smoke is detected and reported by the fire detection and extinguishing (FIDEX) control unit. When the CARGO FIRE PBA is pressed, the IASCs close the CSOVs for the selected cargo compartment.
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21 - Environmental Control 21-23 Ventilation System
AIR PANEL FWD
MAN TEMP
VENT
OFF
CARGO LO HEAT
AFT
OFF
VENT
CARGO FIRE PANEL
HI HEAT
ON
FWD
CARGO BTL
AFT
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
F I RE FIRE E
DMC 1 and DMC 2
IASC 2
AVAIL
F IR E FIRE
IASC 1
FIDEX CONTROL UINIT
FWD Cargo Upstream CSOV
LEGEND CSOV DMC IASC
Cargo Shutoff Valve Data Concentrator Unit Module Cabinet Integrated Air System Controller ARINC 429
FWD Cargo Downstream CSOV
Aft Cargo Downstream CSOV
CS1_CS3_2120_018
rgo Aft Cargo eam Upstream CSOV
Figure 57: Ventilation System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
COMPONENT LOCATION The following component is installed in the ventilation system: •
Cargo shutoff valves (CSOVs)
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21 - Environmental Control 21-23 Ventilation System
Upstream CSOV
Upstream CSOV
Fwd CSOV
CS1_CS3_2120_031
Downstream CSOV
Figure 58: Cargo Shutoff Valves Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
CONTROLS AND INDICATIONS AIR PANEL The forward and aft cargo compartment ventilation is controlled from the AIR panel. The forward CSOVs open anytime the FWD CARGO switch is selected to VENT, LO HEAT, or HI HEAT. The aft CSOVs open when the AFT CARGO switch is selected to VENT. CARGO PANEL If either FIRE PBA is pressed on the CARGO panel, the CSOVs are driven closed by the IASCs.
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21 - Environmental Control 21-23 Ventilation System
AIR COCKPIT
C
FWD
H
CABIN
C
AFT
H
C
FWD
MAN TEMP
CARGO
VENT
OFF
LO HEAT
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
FWD
F IRE IRE L PACK FAIL OFF
FAIL
BTL
AVAIL
AFT
F I RE RE
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
CARGO
FAIL OFF
CARGO PANEL
R BLEED FAIL
APU BLEED
OFF
FAIL CS1_CS3_2120_033
OFF
AIR PANEL
Figure 59: Ventilation System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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INDICATIONS Synoptic Page The FWD CARGO switch selection; OFF, VENT, LO, or HI is displayed on the AIR synoptic page. EICAS Page The forward cargo temperature indication on the engine indication and crew alerting system (EICAS) page is grayed out when the FWD CARGO switch is set to OFF or VENT.
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21 - Environmental Control 21-23 Ventilation System
STATUS FUEL
AIR HYD
DOOR
ELEC
AVIONIC
CKPT 23 °C 22 °C
VENT
15°C
10 °C
CARGO
CLB
FLT CTRL
INFO
FWD 22 °C 22 °C -+ 2
AFT 22 °C 22 °C -+ 2
15 °C
15 °C
CB
73.0
73.3
718
722
EGT
86.5 2410 124 116
R PACK
L PACK
92.2
N1
RAM AIR
TRIM AIR
92.2
86.6 2440 124 116
N2 FF (PPH) OIL TEMP OIL PRESS
TOTAL FUEL (LB) 10905 2735
45 PSI
XBLEED
5435
2735
45 PSI
1400 5.7 ELEV 570
CAB ALT
RATE
ó3
CREW
LDG TEMP (°C)
23
0 OXY 1850
TRIM NU STAB
22
22
ND
4.2 NL RUDDER NR
APU
SYNOPTIC PAGE – AIR
EICAS
CS1_CS3_2123_001
INFO
Figure 60: Ventilation System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
DETAILED DESCRIPTION The cargo shutoff valves are tested at landing by the IASCs, which perform a full open-close cycle to check the motor operation as well as the full close and full open position feedback. Only the in-control channel driver of each IASC is tested during the test. If a failure is detected, switching to the other channel is performed, and the CSOV cycling is commanded again. During normal operation, if the CSOV does not respond to the IASC command, an INFO message is displayed
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
AIR COCKPIT
C
FWD
H
C
CABIN
H
C
FWD
MAN TEMP OFF
VENT
AFT
CARGO LO HEAT
FIDEX CONTROL UNIT
H
FWD
AFT
OFF
PACK FLOW
RECIRC AIR
OFF
HI
OFF
AVAIL
FII R E F
AFT
FII R E F
DMC 1 and DMC 2
IASC 1 TRIM AIR
BTL
VENT
HI HEAT
ON
CARGO
CARGO PANEL OPEN
Position
IASC 2
AIR PANEL
RECIRCULATION FAN
MIX MANIFOLD
Downstream CSOV
Downstream CSOV AFT CARGO COMPARTMENT
LEGEND CSOV DMC FIDEX IASC
From Right Pack
Cargo Shutoff Valve Data Concentrator Unit Module Cabinet Fire Detection and Extinguishing Integrated Air System Controller ARINC 429
Power
FWD CARGO COMPARTMENT
Upstream CSOV
Upstream CSOV
From Trim Air System
CS1_CS3_2120_020
From Left Pack
Figure 61: Ventilation System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the ventilation system:
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_VS_MonitoringTests.fm
21 - Environmental Control 21-23 Ventilation System
CAS MESSAGES Table 23: STATUS Messages MESSAGE
LOGIC
FWD CARGO AIR FWD cargo heating selection selected and confirmed OFF OFF. AFT CARGO AIR OFF
AFT CARGO selected and confirmed OFF.
Table 24: INFO Messages MESSAGE
LOGIC
21 AIR SYSTEM FAULT - FWD CARGO SOV INOP
Upstream of downstream forward cargo SOV failed in position or in unknown position.
21 AIR SYSTEM FAULT - AFT CARGO SOV INOP
Upstream of downstream forward cargo SOV failed in position or in unknown position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
PRACTICAL ASPECTS The cargo shutoff valve (CSOV) has a manual lever used to secure in the closed position before dispatch if the valve has failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
CS1_CS3_2123_003
Manual Shutoff Lever
Figure 62: Cargo Shutoff Valve Practical Aspects Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
21-40 HEATING FLIGHT CREW FLOOR HEATING GENERAL DESCRIPTION The flight crew floor heating system consists of heated mats embedded in each foot rest plate. The heated mats operate as soon as the main DC buses are powered. Thermostats embedded in the mats control the operation of the mats.
COMPONENT LOCATION •
Heated mat
HEATED MAT The heated mats are installed in the foot rest plates in front of each pilot seat.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-144
CS130_21_L2_Heating-FCFH_ComponentLocation.fm
21 - Environmental Control
Pilot Left Heated Mat
Pilot Right Heated Mat
Copilot Left Heated Mat
Copilot Right Heated Mat
CS1_CS3_2140_004
21-40 Heating
Figure 63: Flight Crew Floor Heating General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-145
CS130_21_L2_Heating-FCFH_ComponentLocation.fm
21 - Environmental Control 21-40 Heating
DETAILED DESCRIPTION The heated mats are controlled by thermostats that provide power when the temperature is less than 20°C (68°F) and remove power when the temperature exceeds 40°C (104°F). The control and distribution cabinet 3 (CDC 3) and CDC 4 monitor the current draw. When the current draw exceeds 1 A, the solid-state power controller (SSPC) latches off and reports the overcurrent fault to the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
CDC 3 CDC 3-4-4 L FOOT WARMER DC SSPC 3A BUS 1
HEATED MAT 1 Pilot
Logic: Advanced HEATED MAT 2
CDC 4 CDC 4-4-4 R FOOT WARMER DC SSPC 3A BUS 2
HEATED MAT 3 Co-pilot
Logic: Advanced HEATED MAT 4 CDC 1 and CDC 2
IASC 1 and IASC 2
DMC 1 and DMC 2
OMS
CDC DMC IASC OMS
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Onboard Maintenance System AFDX ARINC 429 TTP BUS
CS1_CS3_2140_005
LEGEND
Figure 64: Flight Crew Floor Heating Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-147
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21 - Environmental Control 21-40 Heating
GALLEY HEATING GENERAL DESCRIPTION The forward and aft galley heating consists of a fan and a heater in a recirculation loop that distributes reheated air next to the service and passenger doors. Air is extracted from the galley area by the galley fan, heated by the galley heater and then exhausted near the floor next to the service and main passenger doors.
Each forward or aft galley heating system is controlled from the flight attendant panels located in the respective forward and aft galley areas. The system status is sent from IASC to the cabin management system (CMS) for fault indication.
The respective IASC monitors the supplemental heating system using a dual-element galley heater temperature sensor (GHTS). The sensor limits the duct temperature to 70°C. Three modes of operation can be selected on a panel installed in the galley area: •
When FAN ON is selected, the fan operates
•
When LO HEAT is selected, the fan operates and the heater operates on a two phase power. With a cabin inlet temperature of 25°C (77°F), the system provides an outlet temperature between 30°C (86°F) and 40°C (104°F)
•
When HI HEAT is selected, the fan operates and the heater operates on three-phase power. With a cabin inlet temperature of 25°C (77°F), the system provides an outlet temperature between 45°C (113°F) and 55°C (131°F)
A dual-temperature sensor is installed at the heater outlet to provide the duct temperature to the IASCs for system monitoring. The forward galley heater receives 3-phase 115 VAC from AC BUS 1 and the aft galley heater receives 3-phase power from AC BUS 2. The forward galley fan is powered by DC BUS 1 and the aft galley fan is powered by DC BUS 2.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
FWD GALLEY PANEL FAN ON
HEATER
IASC 1 CDC 3
HIGH LOW
AC BUS 1 DC BUS 1
ZONE BOX 1
FAN FWD ENTRY WAY
CABIN CONTROLLER
FWD GALLEY
DMC 1 AND DMC 2
HEATER
GALLEY HEAT TEMPERATURE SENSOR CDC 5 AC BUS 2 DC BUS 2
ZONE BOX 2 IASC 2
ON
LEGEND CDC DMC IASC
HEATER HIGH
FAN
LOW
AFT ENTRY WAY
AFT GALLEY PANEL
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 Ethernet
AFT GALLEY
HEATER
GALLEY HEAT TEMPERATURE SENSOR
CS1_CS3_2140_002
FAN
Figure 65: Galley Heating Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
COMPONENT LOCATION The galley heating system consists of: •
Galley fan
•
Galley heater
•
Galley heater temperature sensor
GALLEY FAN The galley fan is attached to the galley structure. GALLEY HEATER The galley heater is attached to the galley structure. GALLEY HEATER TEMPERATURE SENSOR The galley temperature sensor is installed in the duct at outlet of the galley heater.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
Galley Fan
Heater
Temperature Sensor NOTE Fwd galley shown. Aft galley similar.
Outlet
CS1_CS3_2120_014
Outlet
Figure 66: Supplemental Galley Heat Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21 - Environmental Control 21-40 Heating
CONTROLS AND INDICATIONS GALLEY HEAT CONTROLS The galley heaters are controlled from the GALLEY panels located in each galley. Each heater has a fan control and a HIGH or LOW heat selection.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
GA GALLEY ALLEY SHUTOFF SHU UTO OFF LIGHTS
BRT
GALLEY COUNTER
DIM BRT
GALLEY AREA
DIM
FAN
HEATER
HIGH
ON
LOW CHILLER
DEFROST FAIL CS1_CS3_2140_007
ON
Figure 67: Galley Heater Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-153
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21 - Environmental Control 21-40 Heating
GALLEY HEAT INDICATIONS The galley heater faults are displayed on the CMS screen.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control
CS1_CS3_2140_008
21-40 Heating
Figure 68: Galley Heater Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21-155
CS130_21_L2_Heating-GalleyHeating_ControlsIndications.fm
21 - Environmental Control 21-40 Heating
DETAILED DESCRIPTION The heater operation is regulated by a built-in thermostat. For overtemperature protection, thermal fuses in-line with the heating elements open at 183°C (362°F). A heater fault discrete output from the heater assembly is connected directly to the control and distribution cabinet (CDC) that is its power. The galley heater 115 VAC power is removed when a fault occurs. The galley fan remains in operation and only ventilation is provided by the system. A fan fault discrete output signals overcurrent, overvoltage, underspeed (10,000 rpm) or overheat conditions in the fan assembly, which is connected directly to the CDC that provides power to the galley fan. Both the 28 VDC power to the galley fan and 115 VAC power to the galley heater are removed when a fault occurs. In both cases, the fault data is sent to the onboard maintenance system (OMS), and a GALLEY HEATING SYSTEM UNAVAILABLE message is displayed on the cabin management system (CMS) crew terminal. The IASC monitors the galley heater temperature. The galley heater temperature sensor has two elements. One element reports to each of the IASC channels. In case of overheat conditions detected by either element of the galley heater temperature sensor, the heater power is turned off. The IASC commands the CDC to shut off the galley heating when: •
Galley heater outlet temperature reads above 70°C for more than 1 minute
•
The overheat condition is latched until the temperature decreases below 70°C and the GALLEY switch is cycled.
If both elements of the sensor fail, or one element and the opposite IASC channel fail, the heater power is removed. When both elements of the GHTS have failed, the heater is latched off until the GHTS is replaced.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
CDC 3 CDC 3-12-4,5,6 FWD GLY HTR AC SSPC 5A BUS 1
3Ø
Logic: Advanced
FWD GALLEY HEATER
Fwd Galley Heater Fault
FWD GALLEY PANEL FAN
HEATER
CDC 3-6-1 FWD GLY FAN DC BUS 1 SSPC 15A Logic: Advanced
HIGH
ON
FWD GALLEY FAN
Fwd Galley Fan Fault
LOW
ZONE BOX 1
CDC 1
CABIN CONTROLLER
DMC 1 and DMC 2
CDC DMC IASC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 Ethernet TTP
IASC 1 Channel A Channel B
FWD GALLEY HEATER TEMPERATURE SENSOR
NOTE Fwd galley shown. Aft galley similar.
CS1_CS3_2140_003
LEGEND
Figure 69: Supplemental Galley Heat Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21-157
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
21-26 AVIONICS COOLING AND EXTRACTION SYSTEM GENERAL DESCRIPTION The avionics cooling and extraction system consists of three subsystems: •
Air inlet cooling system
•
Air extraction system
•
Backup exhaust and extraction system
In addition, the wing-to-body fairing (WTBF) is ventilated using the air conditioning ram air system. The air inlet cooling system draws air from the galleys, and underfloor area, cools it, and supplies it to the forward and mid avionics bay. The air extraction system draws air from behind the flight deck displays, both forward and mid equipment bays, control and distribution cabinet 5 (CDC 5) and the forward and aft lavatories. The backup exhaust and extraction system provides an alternate means to extract air from behind the flight deck displays, and both forward and mid equipment bays.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
OUTFLOW VALVE
FWD GALLEY
UNDERFLOOR
UNDERFLOOR
AFT GALLEY
CDC 5
M GROUND VALVE AVIONICS FILTER
AVIONICS FILTER
FWD SKIN HEAT EXCHANGER
EXTRACTION FAN 1
EXTRACTION FAN 2
MID BAY FAN
FWD BAY FAN
FWD EQUIPMENT BAY
AVIONICS COALESCENT FILTER
MID SKIN HEAT EXCHANGER
Flight Deck Displays
MID EQUIPMENT BAY
O2 Bottle Compartment
CDC 5
FWD LAV
AFT LAV
FLIGHT DECK
M
FWD AVIONICS EXHAUST VALVE
MID AVIONICS EXHAUST VALVE
M
BACKUP FAN
LEGEND Air Inlet Cooling System Backup Exhaust and Extraction System Air Extraction System
CS1_CS3_2126_012
UNDERFLOOR
Figure 70: Avionics Cooling and Extraction System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-159
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
AIR INLET COOLING SYSTEM GENERAL DESCRIPTION The air inlet cooling system supplies cooled and filtered air into both forward and mid equipment bays in order to lower their ambient operating temperature. There is one air inlet cooling system dedicated to each equipment bay. Each system has a fan, skin heat exchanger, and a filter. The fan extracts air from the galley and underfloor areas. The air is cleaned by a filter before passing through the skin heat exchanger. The skin heat exchanger cools the air using the cold aircraft skin. The cooled air is distributed through piccolo tubes in the equipment bay to cool the equipment. The forward bay fan is powered by DC BUS 2. The mid bay fan is powered by AC BUS 2. The mid bay air inlet cooling system has a ground valve that provides additional air to the mid equipment bay when the air inlet cooling system operates on the ground. The ground valve is powered by DC BUS 2. On the ground, approximately 70% of the cooling comes from the ground valve, and the remaining 30% comes from the galley and control and distribution cabinet (CDC) 5. A coalescent filter is installed downstream of a ground valve, located on the aircraft skin. The coalescent filter extracts the water contained in the outside air and then filters the air for the mid equipment bay. Control of the system is fully automatic. The INLET PBA is normally left in the auto position and only selected OFF when the system has automatically reconfigured to the OFF position. Selecting the INLET PBA OFF turns off the forward and mid bay fans, and closes the ground valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
DMC 1 and DMC 2
IASC 1
CDC 4 CMD/FAULT UNDERFLOOR
IASC 2
CDC 2 AC BUS 2
FWD GALLEY
CDC 5
AFT GALLEY M GROUND VALVE
AVIONICS FILTER
FWD SKIN HEAT EXCHANGER
MID SKIN HEAT EXCHANGER
AVIONICS COALESCENT FILTER
MID BAY FAN
FWD BAY FAN DC BUS 2 FWD EQUIPMENT BAY
MID EQUIPMENT BAY
LEGEND
CDC DMC IASC
Inlet Air Cooling Air Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 Discrete TTP
CS1_CS3_2126_014
AVIONICS FILTER
Figure 71: Air Inlet Cooling Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
COMPONENT LOCATION
MID BAY FAN
The following are major components of the air inlet cooling system:
The mid bay fan is located in the aft cargo compartment on the right side of the fuselage, forward of the cargo door.
•
Forward avionics filter
•
Forward bay fan
AVIONICS COALESCENT FILTER
•
Forward skin heat exchanger
•
Ground valve
The avionics coalescent filter is located in the aft cargo compartment on the right side of the fuselage, forward of the cargo door.
•
Mid bay fan
MID SKIN HEAT EXCHANGER
•
Avionics coalescent filter
•
Mid skin heat exchanger
The aft skin heat exchanger is located in the aft cargo compartment on the right side of the fuselage, aft of the cargo door.
•
Mid avionics filter
MID AVIONICS FILTER
•
Ducting
The mid avionics filter is located in the aft cargo compartment on the right side of the fuselage aft, of the cargo door.
FORWARD AVIONICS FILTER The forward avionics filter is located on the right side of the forward equipment bay. FORWARD BAY FAN The forward bay fan is located on the right side of the forward equipment bay. FORWARD SKIN HEAT EXCHANGER
DUCTING Two separate duct networks provide air to each equipment bay. Each duct runs along the left side of the fuselage. The forward duct connects the forward galleys and underfloor area to the forward equipment bay. The mid duct connects the aft galley and control and distribution cabinet (CDC 5) to the mid equipment bay. The mid bay also has a duct that draws external air when the aircraft is on the ground. The air is distributed in the equipment bays through piccolo tubes.
The forward skin heat exchanger is located on the right side of the forward equipment bay. GROUND VALVE The ground valve is located in the aft cargo compartment on the right side of the fuselage, forward of the cargo door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Galley 4
Mid Skin Heat Exchanger
Ground Valve
Avionics Coalescent Filter Galley 4A (if installed) CDC 5
Mid Bay Fan
Mid Avionics Filter
Ducts
Mid Equipment Bay Piccolo Tubes
Galley 2 Galley 1
Fwd Skin Heat Exchanger
Fwd Bay Fan
CS1_CS3_2126_013
Fwd Avionics Filter
Fwd Equipment Bay Piccolo Tubes
Figure 72: Air Inlet Cooling Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
COMPONENT INFORMATION FORWARD AVIONICS FILTER The forward avionics filter filters air from the forward galleys and underfloor area. The replaceable filter is installed at the inlet to the forward skin heat exchanger. FORWARD SKIN HEAT EXCHANGER The forward skin heat exchanger is installed within the frames of the fuselage structure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Fwd Avionics Filter
CS1_CS3_2126_005
Fwd Skin Heat Exchanger
Figure 73: Forward Avionics Filter and Forward Skin Heat Exchanger Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
GROUND VALVE The GND VLV is installed on the aircraft skin on the lower right side of the fuselage. The ground valve is driven by a linear electrical actuator powered by 28 VDC. A torque limiter protects the actuator motor and prevents inadvertent opening of the valve in flight. The valve position is monitored by microswitches. The valve has a manual override handle on its external side. This handle allows the valve to be closed from outside of the aircraft. AVIONICS COALESCENT FILTER The avionics coalescent filter is located in the duct from the ground valve. The replaceable filter removes contaminants and water from the outside air. A drain line carries the water to the lower fuselage to drain overboard.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Avionics Coalescent Filter GROUND VALVE
Manual Close Lever
CS1_CS3_2126_006
Filter Drain
Figure 74: Ground Valve and Avionics Coalescent Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21-167
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
MID SKIN HEAT EXCHANGER The mid skin heat exchanger is installed within the frames of the fuselage structure. MID AVIONICS FILTER The mid avionics filter filters air drawn from the aft galley and underfloor area. The replaceable filter is installed at the inlet to the mid skin heat exchanger.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Mid Avionics Filter
CS1_CS3_2126_007
Mid Skin Heat Exchanger
Figure 75: Mid Skin Heat Exchanger and Mid Avionics Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
CONTROLS AND INDICATIONS The system is controlled from the EQUIP COOLING panel in the flight deck. In normal operation, the system operation is automatically controlled by the integrated air system controllers (IASCs). The INLET PBA can be selected to turn off the air inlet cooling system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
CARGO
FWD
BTL
AVAIL
FIRE
AFT
FIRE
OVERHEAD PANEL
EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
PRESSURIZATION EMER DEPRESS
CS1_CS3_2126_011
ON
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Figure 76: Avionics Cooling and Extraction System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
DETAILED DESCRIPTION The air inlet cooling system operates all the time on ground and in flight. The air inlet cooling system operation is automatically managed by the IASC main channels, which control the mid ground valve and both forward and mid bay inlet fans through logics embedded in the CDCs. The CDCs drive power supplies, discrete and relay signals located in the electrical distribution system according to requests from the IASCs.
In case a failure is detected on either forward or mid bay fan or the ground valve, an advisory message EQUIP BAY COOL FAULT is displayed. An INFO message provides guidance for any action to be taken before the next aircraft dispatch. The air inlet cooling system can be selected OFF by selecting the INLET PBA on the overhead panel. The aircraft can be dispatched without any operating limitation with the system selected OFF.
In flight, the air inlet cooling system extracts air from the galley, underfloor areas, and CDC 5, cools it through the skin heat exchanger, and then supplies it through a set of distribution lines to cool the avionics bays. On ground, with an outside air temperature (OAT) ranging between 2°C and 33°C, the IASCs command the ground valve open to allow outside air into the system. Before takeoff, the ground valve is driven closed once all doors are signaled closed and locked. If the ground valve fails to close, the valve can be manually closed from outside of the aircraft. After landing, the ground valve is driven open once all doors are not signaled closed. The ground valve is kept closed at landing until the doors open in order to minimize the residual cabin differential pressure, and prevent engine exhaust air going into the filter during the deceleration and taxi phases. In case of smoke detected by the fire detection and extinguishing system (FIDEX) control unit, the IASCs signal the CDCs to remove power from both forward and mid bay fans to automatically shutoff the air inlet cooling system. In case of ditching, when the DITCHING PBA is selected on the PRESSURIZATION panel, the ground valve is automatically driven closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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FWD BAY INLET FAN ON
On/Off
Fault CDC 4
EQUIP COOLING
FWD BAY IN FAN DC BUS 2
EXHAUST VLV AUTO ON ONLY
CDC 2-17-1, 2, 3 MID BAY in AUX AC BUS 2 SSPC 10A
FIDEX CONTROL UNIT
Logic: Always On
PRESSURIZATION EMER DEPRESS
CDC 2-9-15 MID INLET VLV DC SSPC 3A BUS 2
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Logic: Advanced Fault
DITCHING FAIL MAN
LEGEND ON
CDC DMC EPC FIDEX IASC
115 VAC
Logic: Advanced Fault CDC 2-9-9 MID BAY in FAN CMD DC SSPC 3A BUS 2
DMC 1 and DMC 2
AUTO PRESS
28 VDC FWD EQUIPMENT BAY INLET FAN
EPC 2
OFF
ON
25
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electrical Power Center Fire Detection and Extinguishing Integrated Air System Controller AFDX ARINC 429 TTP BUS
MID EQUIPMENT BAY INLET FAN
28 VDC
CDC 2
Position IASC 2
GROUND VALVE
CS1_CS3_2126_019
INLET
Figure 77: Air Inlet Cooling System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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AIR EXTRACTION SYSTEM GENERAL DESCRIPTION The extraction system is composed of two identical extraction fans (EFANs) pulling air from the flight deck displays, both forward and mid equipment bays, CDC 5 enclosure, and the forward and aft lavatories. Extracted air is discharged in closed vicinity of the outflow valve, and then dumped overboard. The extracted airflow is kept below 70°C. The air extraction system operates when the EXHAUST switch on the EQUIP COOLING panel is in the AUTO position. In normal operation, one fan is running, the other is on standby-mode. For reliability and safety reasons, the fans are alternately operated on daily basis. EFAN 1 is powered by AC BUS 2. The motor controller is powered by DC BUS 2. EFAN 2 is powered by AC BUS 1. The motor controller is powered by DC BUS 1. On the ground, the EFAN operates at maximum speed to maximize heat extraction, and maintain the equipment bay temperature as low as possible. In flight, the EFAN operates at a lower speed to maintain extraction flow below the outflow valve (OFV) exhaust flow value.
NOTE On the ground, powering the aircraft with batteries only must be limited to 5 minutes. After 5 minutes, AC power must be provided for EFAN operation to ensure adequate equipment bay cooling.
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EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
Safety CH
DMC 1 and DMC 2
IASC 1 CH A
Safety CH IASC 2 CH A
FLIGHT DECK Flight Deck Displays
O2 Bottle Compartment
FWD LAV
AFT LAV
FWD EQUIPMENT BAY
MID EQUIPMENT BAY
DC BUS 2
CDC 5
EXTRACTION FAN 1
AC BUS 2
CDC DMC IASC
CDC 2
Exhaust Air DC BUS 1 Control and Distribution Cabinet Data Concentrator Unit Module Cabinet AC BUS 1 Integrated Air System Controller ARINC 429 CDC 1 Discrete
OUTFLOW VALVE
EXTRACTION FAN 2
CS1_CS3_2126_016
LEGEND
Figure 78: Air Extraction System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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COMPONENT LOCATION The following is a major component of the main air extraction system •
Extraction fans
•
Ducting
EXTRACTION FANS The extraction fans are located in the environmental control system bay. DUCTING The extraction ducting runs along both side of the fuselage, connecting the flight deck display lines, oxygen compartment, and lavatories. Piccolo tubes draw air from the forward and mid equipment bays. The ducts join in a common manifold in the environmental equipment bay. The manifold connects all of the ducting to the extraction fans. The extraction fan exhaust discharges the exhaust air into the environmental bay near the outflow valve.
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Control and Distribution Cabinet 5
Mid Equipment Bay Piccolo Tubes
Lavatory E Extraction Fans
Fwd Equipment Bay Piccolo Tubes Flight Displays and Center Console
Extraction Fan Exhaust
CS1_CS3_2120_004
Ducts
Lavatory A Oxygen Compartment
Figure 79: Main Air Extraction System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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DETAILED DESCRIPTION In normal operation, one extraction fan (EFAN) is running, and the other extraction fan is on standby mode. Both avionics exhaust valves (AEV) are closed and the backup fan is off. The EFAN selection and the system reconfiguration are controlled by the IASC safety channels. Each safety channel controls the power supply of one EFAN by commanding the relay coil that switches the 115 VAC to the EFAN. Each EFAN is operated on different days. Automatic switching occurs on the ground. The fan selection logic is done by the IASCs safety channel, based on the date provided by the data concentrator unit module cabinet (DMC). Each extraction fan provides low-speed and internal fault monitoring feedback to the IASCs through a CAN BUS for troubleshooting and monitoring of the EFANs. A fault discrete signal is also sent from the EFAN to the IASC safety channels for monitoring of the EFAN internal failures. Fault detection is done by the fan motor control. Fan speed, overheat, overcurrent, and overvoltage faults are monitored. If one EFAN fails, the system automatically switches to the other EFAN, and an EQUIP BAY COOL FAULT advisory message is displayed. On the ground, the EFAN operates at full speed in order to maximize heat extraction, and maintain equipment bay ambient temperature as low as possible. In flight, the EFAN operates at a lower speed to maintain extraction flow below the outflow valve exhaust flow, and to ensure no equipment bay airflow is recirculated into the distribution system. Dispatch is possible with one EFAN failure, if the backup exhaust system is operative.
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EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
DMC 1 and DMC 2
IASC 1 EFAN 1 Fault
EFAN 1
EXTR Fan CTRL DC BUS 2
5
20 EXTR Fan 1 AC BUS 2
20 20
External FAN 1 Discrete CDC 1
NOTE EFAN 1 shown. EFAN 2 similar.
EPC 2
LEGEND Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Extraction Fan Electrical Power Center Integrated Air System Controller AFDX ARINC 429 CAN BUS
CS1_CS3_2126_017
CDC DMC EFAN EPC IASC
Figure 80: Main Air Extraction System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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BACKUP EXHAUST AND EXTRACTION SYSTEM GENERAL DESCRIPTION A backup cooling system is provided for the flight deck displays and avionics cooling functions with the addition of two duct networks, extracting air from behind the flight deck displays, and from the forward equipment bay and the mid equipment bay. Each duct connects to one avionics exhaust valve (AEV) installed on the aircraft skin. The AEV is a shutoff valve, which opens and exhausts the airflow overboard. Each AEV is powered and monitored by both IASCs. The amount of airflow extracted is mainly dependent on the cabin to ambient pressure differential existing between the inlet and the outlet of a flow restrictor integral to the aircraft skin. A third backup extraction fan is fitted on top of the mid avionics bay enclosure, and provides air extraction through an opening in the center ceiling of the mid avionics bay. The backup fan is powered by DC ESS BUS 1.
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EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
DMC 1 and DMC 2
IASC 2 CH A
CH B
IASC 1 CH A
CH B
FLIGHT DECK O2 Bottle Compartment
DC ESS BUS 1 BACKUP FAN
VENTS
MID EQUIPMENT BAY
FWD EQUIPMENT BAY
M
FWD AVIONICS EXHAUST VALVE
VENTS
M
MID AVIONICS EXHAUST VALVE
LEGEND DMC IASC VENTS
Exhaust Air Data Concentrator Unit Module Cabinet Integrated Air System Controller Ventilated Temperature Sensor ARINC 429 Discrete
CS1_CS3_2126_015
Flight Deck Displays
Figure 81: Backup Exhaust and Air Extraction System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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COMPONENT LOCATION The following are major components of the backup exhaust and air extraction system: •
Avionics exhaust valves
•
Backup fan
•
Ducting
•
Avionics bay ventilated temperature sensors (Refer to figure 83)
AVIONICS EXHAUST VALVES The forward avionics exhaust valve is located in the forward cargo compartment on the right side of the lower fuselage, forward of the cargo door. The aft avionics exhaust valve is located in the aft cargo compartment on the right side of the lower fuselage, forward of the cargo door. BACKUP FAN The backup fan is mounted on the ceiling of the mid equipment bay. DUCTING There are two duct networks for the backup exhaust and air extraction system. The forward duct runs along the left side of the aircraft, connecting the flight deck displays, and forward equipment bay to the forward avionics exhaust valve. The mid equipment bay ducting connects to the mid avionics exhaust valve. Air from the flight deck displays, forward, and mid equipment bays is drawn through piccolo tubes, and exhausted overboard when the avionics exhaust valves are open.
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Mid Avionics Exhaust Valve Backup Fan
Fwd Avionics Exhaust Valve Ducts Flight Deck Display Piccolo Tubes
CS1_CS3_2120_005
Ducts
Mid Equipment Bay Piccolo Tubes
Fwd Equipment Bay Piccolo Tubes
Figure 82: Backup Exhaust and Air Extraction System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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AVIONICS BAY VENTILATED TEMPERATURE SENSORS An avionics ventilated temperature sensor (AV-VENTS) is installed in each of the forward and mid equipment bay ceilings.
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Note
Fwd equipment bay
Fwd equipment bay shown. Mid equipment bay similar.
CS1_CS3_2126_010
Ventilated Temperature Sensor
Figure 83: Avionics Bay Ventilated Temperature Sensor Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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DETAILED DESCRIPTION The backup exhaust and extraction systems are in standby during normal operation. Each bay has an avionics ventilated temperature sensor (AV-VENTS) sensor for monitoring. The AV-VENTS sensor has two temperature elements with one reporting to each IASC. A small fan draws air across the elements. If the temperature exceeds 67°C (153°F) for more than 30 seconds, an EQUIP BAY OVHT warning message is displayed, and the auxiliary power unit (APU) (multifunction) horn sounds. The message is cleared when the temperature decreases to 62°C (144°F), and the exhaust switch on the EQUIP COOLING panel is moved from the AUTO position.
If the DITCHING PBA is selected on the PRESSURIZATION panel, the AEVs are automatically driven closed to prevent water from entering the aircraft. Both AEVs and the backup fan are tested 3 minutes after landing to confirm the backup exhaust and extraction systems is operative. In case a failure is detected on either AEV or on the backup fan at landing, an EICAS EQUIP BAY COOL FAULT advisory and relevant INFO messages are displayed. One or both AEVs can be failed for dispatch. Aircraft dispatch is not possible with the backup fan inoperative.
If both EFANs fail, or both channels of one AV-VENTS fail, the IASCs automatically open both avionics exhaust valves (AEVs), starts the backup fan, and an EICAS EQUIP BAY COOL FAIL caution message is displayed. The EXHAUST switch on the EQUIP COOLING panel is set to ON to match the new system configuration. In the event of smoke detected in the avionics bay, both AEVs are automatically driven open and the backup fan is kept off by the IASCs if no overheat is detected. The EXHAUST switch is moved to VLV ONLY to match the system configuration. A hardwired contact on the EXHAUST switch provides additional control of the fan.
NOTE When the EXHAUST switch is set to ON, a discrete hardwired signal from the switch turns on the backup fan. On ground when operating on the batteries only, as well as during approach when the ram air turbine (RAT) generator switches off, the electrical distribution system disables the backup fan in order to protect the batteries.
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ECS Safety Sensor DC ESS BUS 3
LEGEND CTL DMC EPC FIDEX IASC TEMP TLC VENTS
Control Data Concentrator Unit Module Cabinet Electrical Power Center Fire Detection and Extinguishing Integrated Air System Controller Temperature TRU Line Contactor Ventilated Temperature Sensor ARINC 429
3
RIGHT CIRCUIT BREAKER PANEL
Fwd Temp Mid Temp
EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
Backup Fan Ctl
Cmd
FWD AVIONICS EXHAUST VALVE
IASC 1
FIDEX CONTROL UNIT
DMC 1 and DMC 2
Open/Close Position
OFF
Mid Bay Altn Fan DC ESS BUS 1
FWD EQUIPMENT BAY VENTS
Fault 20
28 VDC
EPC 1 TLC 3 Aux Fan Ctl In TLC 3 Aux Fan Ctl Out EPC 3
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Backup Fan Ctl
DITCHING
MAN
Fwd Temp
Cmd
AUTO PRESS FAIL
Mid Temp MID EQUIPMENT BAY VENTS Open/Close Position
ON
MID AVIONICS EXHAUST VALVE IASC 2
CS1_CS3_2126_018
PRESSURIZATION
Fan on/off BACKUP FAN
Figure 84: Backup Exhaust and Air Extraction System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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COMPONENT INFORMATION AVIONICS EXHAUST VALVES The avionics exhaust valve (AEV) is an electrically actuated butterfly valve powered by a 28 VDC motor. The valve remains in its last position if power is lost. Microswitches provide position feedback to the IASCs. The AEV has a manual lever which can used to secure the valve in the open position after a failure.
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Manual Lever
Adapter
CS1_CS3_2126_009
Screen
Figure 85: Avionics Exhaust Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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BACKUP FAN The backup fan is mounted on the ceiling of the mid equipment bay. The fan runs at a single-speed when powered. A check valve closes off the fan exhaust when the fan is not running. The backup fan deflector is installed at the backup fan exit to direct smoke away from the floor sills where it could enter the passenger cabin.
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Check Valve
Mid Equipment Bay Ceiling
Backup Fan Deflector
CS1_CS3_2126_008
Backup Fan
Figure 86: Backup Fan Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CONTROLS AND INDICATIONS The system is controlled from the EQUIP COOLING panel in the flight deck. In normal operation, the EXHAUST rotary switch is in the AUTO position and the system operation is automatically controlled by the IASCs. If a fault occurs in the main extraction system, the IASC automatically selects the backup exhaust and extraction system. The EXHAUST rotary switch is moved to the ON position to match the system configuration. If smoke is detected in an avionics bay, the backup extraction fan can be selected off using the VLV ONLY rotary switch position.
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CARGO
FWD
BTL
AVAIL
FIRE
AFT
FIRE
EQUIP COOLING OVERHEAD
INLET
EXHAUST VLV AUTO ON ONLY
OFF
PRESSURIZATION EMER DEPRESS
CS1_CS3_2126_020
ON
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Figure 87: Avionics Cooling and Extraction System - Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the avionics cooling and extraction system:
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CAS MESSAGES
Table 28: STATUS Messages Table 25: WARNING Message
MESSAGE EQUIP BAY OVHT
MESSAGE INLET AIR OFF
LOGIC Fwd or mid avionics bay overtemperature above 67° C.
EQUIP BAY COOL FAIL
MESSAGE
LOGIC Loss of both extraction FANS confirmed by either IASC control or safety channel. Both FANS are running. Total loss of one zone temperature measurement, with electrical power supplied, and with either IASC safety channel operational.
Table 27: ADVISORY Messages MESSAGE EQUIP BAY COOL FAULT
LOGIC FWD or MID equipment bay cooling fault
Table 28: STATUS Messages MESSAGE
LOGIC
EXHAUST AIR ON
Exhaust switch selected to ON. Supplemental fan turns on and valves open.
EXHAUST AIR VLV ONLY
EXHAUST selector to VLV ONLY and both AEV commanded OPEN.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
INLET PBA selected OFF and both forward and mid inlet fans commanded OFF and the mid ground valve commanded closed.
Table 29: INFO Messages
Table 26: CAUTION Message MESSAGE
LOGIC
LOGIC
21 EQUIP BAY COOL FAULT EFAN INOP
Extraction FAN 1 or FAN 2 inoperative.
21 EQUIP BAY COOL FAULT EFAN CAN BUS INOP
Extraction FAN 1 or FAN 2 CAN BUS inoperative.
21 EQUIP BAY COOL FAULT ALTN FAN INOP
Backup FAN inoperative.
21 EQUIP BAY COOL FAULT SUPP FAN INOP
Supplementary bay fan inoperative.
21 EQUIP BAY COOL FAULT IFAN INOP
FWD or MID Inlet FAN inoperative.
21 EQUIP BAY One channel of either FWD or MID or SUPP bay ventilated sensor failed. COOL FAULT AVIO TEMP SNSR REDUND LOSS 21 EQUIP BAY COOL FAULT FWD AVIO EXHAUST VLV INOP
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FWD AEV failed in position or unknown position.
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Page Intentionally Left Blank
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Table 29: INFO Messages MESSAGE
LOGIC
21 EQUIP BAY COOL FAULT MID AVIO EXHAUST VLV INOP
MID AEV failed in position or unknown position.
21 EQUIP BAY COOL FAULT MID GND VLV INOP
MID equipment bay ground valve failed in position.
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WING-TO-BODY FAIRING VENTILATION SYSTEM GENERAL DESCRIPTION The wing-to-body fairing (WTBF) ventilation system provides ventilation and heat extraction in the forward section of the body fairing to limit the temperature of the WTBF composite structure. The WTBF ventilation system induces secondary airflow to extract heat radiated by the packs and the high-pressure bleed air ducting. Each air conditioning unit (ACU) ram air exhaust has a low-pressure ejector and mixer duct to induce ventilation. The ejector and mixer duct are located just downstream the ACU plenum exhaust and ram air regulating valve (RARV). When the packs are operating, the heat exchanger cooling air is discharged through the RARV into the ram air exhaust. This airflow draws the WTBF compartment air into the ram air exhaust to be discharged overboard.
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CS1_CS3_2150_003
Ram Air Regulating Valve
Figure 88: Wing-to-Body Fairing Ventilation System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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HIGH-PRESSURE PNEUMATIC DUCT AREA COOLING Ram air is used to cool the area around the high-pressure pneumatic duct. Air is tapped off the ram air inlet ducts at the heat exchanger inlet. Two supply tubes feed a piccolo tube that direct air towards the bulkhead area to ensure that components installed in that area are not overheated.
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Piccolo Tube
High-Pressure Pneumatic Duct
Ram Air Duct
Air Cycle Machine Heat Exchanger Inlet
CS1 CS3 2150 031 CS1_CS3_2150_031
Supply Tubes
Figure 89: High-Pressure Pneumatic Duct Area Cooling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-30 CABIN PRESSURE CONTROL SYSTEM GENERAL DESCRIPTION
MANUAL MODE
The cabin pressure control system (CPCS) automatically controls the pressure inside the cabin to a safe and comfortable level by modulating the flow of pressurized air through the outflow valve (OFV).
The manual mode is selected using the AUTO PRESS PBA. IASC channel B controls the outflow valve through the safety channel, and the MAN RATE rotary switch adjusts the cabin rate of change. In this mode, both automatic channels are inactive. Only one manual channel is in control at a time.
The system consists of two integrated air system controllers (IASCs). Each IASC has three channels: channel A, channel B, and a safety channel. The pressurization system has three independent modes of operation each driving its own actuator on the outflow valve through the CABIN PRESSURE relay. There are two automatic modes controlled by IASC channel A and one manual mode operated from the PRESSURIZATION panel through the IASC safety channel. The manual mode also provides overpressure protection. Auto channel 1 is powered by DC ESS BUS 1, and auto channel 2 is powered by DC ESS BUS 2. The manual channel is powered by the IASC in control. At takeoff, the outflow valve starts to close to pre-pressurize the cabin. During flight, the OFV is normally slightly open in order to maintain the cabin pressure. After landing, the OFV is driven full open to depressurize the aircraft. On the ground the OFV is normally fully open to equalize the inside pressure to the outside ambient pressure to ensure the doors can be opened safely. AUTOMATIC MODE In normal mode, the system operates automatically during all flight phases based on predetermined schedules using IASC channel A. The cabin altitude reference is based on inputs from aircraft systems through the data concentrator unit module cabinets (DMCs). Both IASCs are active in automatic mode, but only one is in control. IASC 1 is in control on odd days, and IASC 2 is in control on even days. If one IASC fails in auto mode, the second IASC automatically takes over. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
DITCHING The DITCHING function is always available, whether the system is in AUTO mode or in MANUAL mode. When the DITCHING PBA is selected, the OFV is driven open to depressurize the aircraft. The packs are shut OFF automatically, and all valves below the flotation line are closed. Prior to ditching, the OFV is driven close to avoid water intake. EMERGENCY DEPRESSURIZATION Emergency depressurization is always available in AUTO or MANUAL mode. When the EMER DEPRESS PBA is selected, the OFV is driven open. The rapid depressurization function has protection against excessive cabin rate of change, and cabin altitude. SAFETY FUNCTIONS Two safety valves (SFVs) protect the pressurized fuselage against overpressure conditions. The SFV open if the cabin differential pressure begins to exceed the structural limits. The SFVs also provide negative pressure-relief. A negative pressure-relief valve (NPRV) provides additional protection against negative differential pressure. Pressure equalization valves (PEV) maintain pressure equilibrium between the pressurized cabin, and the cargo compartment. The large PEV allows the air from cabin into the cargo compartment, while the small PEV allows the airflow out of the cargo compartment into the cabin.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
DITCHING AUTO PRESS FAIL
DMC 1 and DMC 2 IASC 1
ON
MAN
MAN RATE
PAX OXY
CDC 1 DPLY DN
Auto 1
DC ESS BUS 1
UP
PRESSURIZATION PANEL
CABIN PRESSURE RELAY
CDC 2 DC ESS BUS 2
Manual
Auto 2 OUTFLOW VALVE
IASC 2 LEGEND Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Negative Pressure-Relief Valve Pressure Equalization Valve – Large Pressure Equalization Valve – Small Safety Valve AFDX ARINC 429 RS 422
PEV L FWD CARGO COMPARTMENT
SFV 2 PEV S PEV L
AFT CARGO COMPARTMENT
PEV S
NPRV AFT Pressure Bulkhead
CS1_CS3_2130_017
CDC DMC IASC NPRV PEV L PEV S SFV
SFV 1
Figure 90: Cabin Pressure Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
TYPICAL FLIGHT PROFILE
Climb
In normal automatic cabin pressure control mode, a typical flight profile includes the following sequences:
The climb sequence starts 6000 ft above takeoff altitude or 10 minutes after takeoff. The cabin rate of climb is controlled for passenger comfort. The cabin climb profile varies directly with the aircraft climb rate.
•
Prepressurization
•
Takeoff
Cruise
•
Return to base
•
Climb
•
Cruise
The cruise sequence starts when the aircraft rate of climb is less than 500 fpm for 10 seconds. The cabin altitude is maintained in accordance with the pressurization schedule. The cabin altitude is controlled to a maximum of 7840 ft at 41000 ft aircraft altitude.
•
Descent
•
Depressurization
Descent
Prepressurization Prepressurization starts when the aircraft is on the ground and the thrust levers are advanced to the takeoff range. The aircraft pressurizes to 300 ft below the takeoff field elevation to minimize the pressure bump at takeoff. Takeoff When the aircraft leaves the ground, the takeoff sequence is initiated. The takeoff sequence is maintained until the aircraft is more than 6,000 ft above the takeoff field elevation or 10 minutes has elapsed. The main function of the takeoff sequence is to avoid having to reselect the landing altitude in case of a need to return and land at the takeoff field.
The descent sequence starts when the aircraft descends at more than 500 fpm for 10 seconds. Cabin altitude descends following the pressurization schedule until the cabin altitude is 300 ft below the landing field elevation. This minimizes the pressure bump at landing. Depressurization Mode The depressurization sequence starts when the aircraft lands. The cabin pressurizes to the field elevation at a rate of 300 fpm for 45 seconds, and then is driven fully open to minimize any pressure differential that might prevent the doors from opening safely.
Return-to-Base The return-to-base sequence starts if the aircraft descends at more than 500 fpm for more than 10 seconds during the takeoff sequence. The aircraft depressurizes to 300 ft below the landing field elevation.
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21 - Environmental Control 21-30 Cabin Pressure Control System
rc
ra
ft
Al
tit
ud e
41,000 ft
Ca
bi
n
Al
tit
ud
e
Ai
7840 ft
500 ft/min
Takeoff Elevation
300 ft/min
Landing Elevation -300 ft
CS1_CS3_2130_011
-300 ft
Figure 91: Typical Flight Profile Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
COMPONENT LOCATION
PRESSURE EQUALIZATION VALVES
The following components are part of the cabin pressure control system:
The pressure equalization valves (PEVs) are located in the cargo compartment entrance way. The forward cargo compartment PEVs are aft of the cargo door. The aft cargo compartment PEVs are forward of the cargo door.
•
Outflow valve (OFV)
•
Water activated check valve
•
Muffler
•
Safety valve (SFV)
•
Negative pressure-relief valve (NPRV)
•
Pressure equalization valves (PEVs)
OUTFLOW VALVE The outflow valve is located in the environmental control system bay on the bulkhead. WATER ACTIVATED CHECK VALVE The water activated check valve is located in the wing-to-body-fairing (WTBF). MUFFLER The muffler is located in the WTBF. SAFETY VALVE There are two safety valves (SFVs) located on the aft pressure bulkhead. NEGATIVE PRESSURE-RELIEF VALVE The negative pressure-relief valve (NPRV) is located on the aft pressure bulkhead.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Safety Valve Pressure Equalization Valve – Large Pressure Equalization Valve – Small Water Activated Check Valve
Negative Pressure-Relief Valve
CS1_CS3_2130_013
Pressure Equalization Valve – Large Pressure Equalization Valve – Small
Muffler Outflow Valve
Figure 92: Pressurization Components Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
COMPONENT INFORMATION OUTFLOW VALVE The outflow valve modulates the discharge airflow to control the cabin pressure in both AUTO modes and in MANUAL mode. The OFV actuator assembly consists of two identical auto channels and a manual channel. A travel limiter physically limits the outflow valve opening when deployed. When the aircraft altitude exceeds 17,000 ft, the travel limiter pin extends to limit the outflow valve travel.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Outflow Valve Actuator
CS1_CS3_2130_014
Travel Limiter Pin Travel Limiter
Figure 93: Outflow Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
WATER ACTIVATED CHECK VALVE The water activated check valve is a mechanical flapper installed on the environmental control system bay bulkhead. The flapper is supported by a bracket assembly under normal conditions. If the aircraft ditches, the rising water level closes the flapper. MUFFLER The optional muffler reduces the outflow valve discharge air noise level.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Bracket Assembly
Water Activated Check Valve Pressure Bulkhead
Muffler
CS1_CS3_2130_007 S1 CS3 2130 007
Water Activated Check Valve
Figure 94: Outflow Valve, Water Activated Check Valve, and Muffler Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
SAFETY VALVE Two safety valves protect against any overpressure condition. The safety valves are pneumatically-actuated on the ground or in flight. The SFVs provide protection for both overpressure relief, and negative pressure-relief. NEGATIVE PRESSURE-RELIEF VALVE The negative pressure-relief valve (NPRV) is a mechanical check valve with spring-loaded flappers. The NPRV is acted on by ambient pressure on one side, and cabin pressure on the other side to sense the differential pressure.
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21 - Environmental Control 21-30 Cabin Pressure Control System
A
B
B NEGATIVE PRESSURE-RELIEF VALVE
A SAFETY VALVE AFT PRESSURE BULKHEAD
CS1_CS3_2130_009
Flappers
Figure 95: Safety Valves and Negative Pressure-Relief Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURE EQUALIZATION VALVES The system maintains pressure equalization between the cabin and forward and aft cargo using four pressure equalization valves (PEVs). The PEV is a fully mechanical device that prevents excessive pressure between the cargo compartment and cabin underfloor in both directions (inflow and outflow). Two PEVs are installed on each cargo compartment entryway. Each cargo compartment has one large cargo PEV, and one small PEV. The large PEV allows the air from the cabin into the cargo compartment, and the small PEV allows the airflow out of the cargo compartment into the cabin. The large PEV prevents the cabin from pressurizing when the cargo doors are not safely closed, and prevents the cargo blowout panels from opening. The small PEV allows cargo compartment pressure to equalize pressure with cabin pressure during fast climb.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Pressure Equalization Valve – Large
NOTE Aft cargo shown. Fwd cargo similar.
CS1_CS3_2130_008
Pressure Equalization Valve – Small
Figure 96: Pressure Equalization Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
DETAILED COMPONENT INFORMATION INTEGRATED AIR SYSTEM CONTROLLER The system consists of two identical integrated air system controllers (IASCs). Each IASC incorporates channel A and channel B and a safety channel. A cabin pressure sensor is fitted to each channel A and channel B of the IASC so that the cabin pressure control system receives four independent cabin pressure signals. The safety channel receives cabin pressure information from the channel B sensor. The safety channel limits cabin altitude and cabin rate of change, and sends cabin altitude and cabin pressure rate of change to the data concentrator unit module cabinet (DMC) by an ARINC 429 BUS.
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21 - Environmental Control 21-30 Cabin Pressure Control System
IASC 1
Channel A
Channel B
Safety Channel
• Automatic pressure control
• Manual pressure control
• Cabin altitude limitation
• Emergency depressurization mode
• Emergency depress mode
• Ram air (OFV forced open)
• Ditching sequence
• Ditching sequence
• Manual pressure control
PS
• Cabin altitude limitation
PS
• Cabin pressure sensor
Data Exchange
DMC
IASC 2
Channel B
Safety Channel
• Automatic pressure control
• Manual pressure control
• Cabin altitude limitation
• Emergency depressurization mode
• Emergency depress mode
• Ram air (OFV forced open)
• Ditching sequence
• Ditching sequence
• Manual pressure control
PS
• Cabin altitude limitation
PS
ARINC 429 DMC Data Concentrator Unit Module Cabinet OFV Outflow Valve PS
Pressure Sensor
• Cabin pressure sensor CS1_CS3_2190_005
LEGEND
Channel A
Figure 97: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
CONTROLS AND INDICATIONS PRESSURIZATION PANEL The PRESSURIZATION panel has controls for: •
Automatic cabin pressure control
•
Manual cabin pressure control
•
Rapid cabin pressure-relief (EMER DEPRESS)
•
Ditching
The system is normally in AUTO mode for cabin pressurization. If the cabin pressure AUTO fails, the amber FAIL light illuminates. Selecting the AUTO PRESS PBA to MAN, switches the cabin pressure control mode to MANUAL. In manual mode, the MAN RATE knob adjusts the cabin altitude. Counterclockwise selection decreases the cabin altitude at a maximum rate of -2,500 ft/min. Clockwise selection increases the cabin altitude at a maximum rate of +2,500 ft/min. There are detents at the +1,000 ft positions on the switch. Selecting the guarded EMER DEPRESS PBA to ON, opens the outflow valve. The cabin altitude and rate limitations functions (15,000 ft) remain active. The white OPEN legend indicates the outflow valve is being forced open. The DITCHING PBA closes all of the valves mounted on the aircraft located below the flotation line, in preparation for a ditching. Selecting the guarded DITCHING PBA opens the outflow valve when altitude is below 15,000 ft. All valves are closed electrically. Prior to ditching, the outflow valve is driven closed.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
DITCHING AUTO PRESS FAIL
ON
MAN
MAN RATE
PAX OXY
DPLY UP
PRESSURIZATION PANEL
CS1_CS3_2130_004
DN
Figure 98: Pressurization Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
FMS PAGE The landing elevation is selected via the flight management system (FMS) through an automatic mode or two backup manual modes. When the flight plan is programmed in the FMS ROUTE tab, the landing elevation of arrival airport data is displayed and also sent via an ARINC 429 BUS to the IASCs. The landing elevation is also set using: •
Control tuning panel (CTP)
•
AVIONIC CTP page
Using the CTP, the LDG ELEV left line select key switches between FMS and MAN, and the LDG ELEV right line select key changes the MAN elevation. The CTP input can be overridden using the avionics virtual CTP FMS/MAN LDG ELEV selection to set the landing elevation manually. In either manual mode, the landing elevation is sent to the IASCs via an ARINC 429 BUS.
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21 - Environmental Control 21-30 Cabin Pressure Control System
BRT
HDG
MAG/TRUE
OFF
TUNE/ MENU
11234
AIR
DOOR
ELEC
FLT CTR CTRL TRL L
FUEL FUE EL
HYD YD
AVIONIC
INFO
CB
AVIOO
IDENT
BRG 1/2>
LDG ELEV FMS/MAN
STATU STATUS TATUS S
L CTP OVRD 850 FT
TUNE/DATA
NAV SRC
FMS 1 BARO
ACT FPLN 24R
NAV SRC
000
FMS 1 BARO
UNITS
FINAL APPR CRS
108.50
251T
RWY THRESHOLD ELEV 126 FT
GS ANGLE
3.00°
MINIMUMS
UNITS
MINIMUMS
OFF BRG 2
BRG 1
BRG 2
OFF
OFF
OFF
OFF
HDG
FPV CAGE
HDG
FPV CAGE
MAG TRUE
ON OFF
MAG TRUE
ON OFF
FMS MAN
YES
000
BRG 1
LDG ELEV
WGS-84
CRS
00.00 STD IN
OFF
FREQUENCY
ARRIVALS...
CRS
00.00 STD IN
CONTROL TUNING PANEL (CTP)
Landing Elevation Normally Set on Arrival Page in FMS
R CTP OVRD
1/2
RETURN
ARRIVAL DATA KLAX ILS
CTP TP
850
FT
LDG ELEV
FMS MAN
DONE DON NE SYNOPTIC PAGE – AVIONICS
FMS ARRIVAL PAGE
000
FT CS1_CS3_2130_005
MINIMUMS BARO/RAD
Figure 99: FMS, Control Tuning Panel, and AVIONIC CTP Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
INDICATIONS The cabin pressure control system information is displayed on the engine indication and crew alerting system (EICAS) page or the AIR synoptic page. The CAB ALT value is the highest recorded value by any of the six IASC channels. If the landing elevation is set manually, MAN appears next to the LDG ELEV value. An arrow is displayed beside the RATE to indicate the direction of the cabin rate of climb. •
If the cabin rate of climb increases, an upward white arrow appears
•
If the cabin rate of climb decreases, a downward white arrow appears.
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21 - Environmental Control 21-30 Cabin Pressure Control System
TOTAL FUEL (KG) 10905 2735
1
5435
5500 8.0 ELEV 560
2735
CAB ALT
RATE
ó3
CREW
LDG 7(03 (°C)
23
100 OXY 2000
TRIM NU STAB
LO
22
22
CAB ALT Description
Value
4.2
ND
NL RUDDER NR
INFO
XXXXX
Cabin altitude above warning threshold
XXXX or XXXXX
Cabin altitude above caution threshold
XXXXX
Cabin altitude within normal range
------
Invalid value
Value
ǻ3 Description
Value
36, 36,
XXXXX
Cabin altitude manually selected or pressurization system in manual mode
XXXXX
Cabin altitude actual value
-----
Invalid value
3
(,&$63$*(
5 -45
RATE Description
X.X
Delta P above warning threshold
XX.X
Delta P within normal range
--.-
Invalid value
Value
560
5 -45
XBLEED
36,
560 MAN -----
LDG ELEV Description LDG ELEV selected from FMS LDG ELEV selected through CTP or AVIONIC synoptic page Invalid value
$38
ó3 LDG ELEV
5500 8.0 560
6<1237,&3$*(±$,5
RATE CREW OXY
100 2000
1
Air block not shown in EICAS compressed mode.
2
Air block only shown on SYNOPTIC page when EICAS in compressed mode.
3
The cabin rate arrow and digital readout in cyan for 3 seconds when manually turns selected.
2
CS1_CS3_2130_006
NOTE
CAB ALT
Figure 100: EICAS and AIR Synoptic Page Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
DETAILED DESCRIPTION AUTO/MANUAL SELECTION The system is normally in AUTO mode for cabin pressurization. If the IASC controlling the automatic pressurization fails, the standby channel automatically takes over. If both IASC automatic channels fail, the AUTO PRESS PBA indicates FAIL, and an AUTO PRESS FAIL caution message is displayed. When the AUTO PRESS PBA is selected to MAN, the cabin pressure control mode switches to manual. The AUTO PRESS PBA MAN indication illuminates white and a CABIN PRESS MAN status message is displayed. The AUTO PRESS PBA can also be used to select active channels if required. When in automatic mode, pressing the AUTO PRESS PBA selects the manual channel of the IASC. A second press of the pushbutton annunciator (PBA) selects the other IASC automatic channel. A third press of the PBA selects the other IASC manual channel.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION
Auto 1 IASC 1
EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
Manual IASC 1
AUTO PRESS FAIL MAN
Auto 2 IASC 2
MAN RATE
DPLY DN
UP P
Manual IASC 2
CS1_CS3_2130_003
PRESSURIZATION PANEL
Figure 101: Cabin Pressure Control System Selection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
INTERFACE The outflow valve (OFV) actuator assembly consists of two identical auto channels and a manual channel. The auto channels of the OFV communicate with their corresponding integrated air system controller (IASC) channel A via an RS 422 BUS. The manual channel is hardwired to the IASCs through the CABIN PRESSURE relay, located in electrical power center 2 (EPC 2). The CABIN PRESSURE relay allows both IASCs to manually command the OFV. Only one controller at a time is connected to the MANUAL channel of the OFV. The outflow valve commands are provided by the IASC in control. The outflow valve manual channel provides position feedback to each IASC.
The aircraft cannot be dispatched if the travel limiter pin is failed closed. If the pin fails in the extended position, the aircraft can be dispatched but is limited to a maximum altitude of 25,000 ft. Cabin Altitude Limitation The cabin altitude limitation system limits cabin altitude to a maximum of 15 000 ft. When the altitude limit is reached, a signal is sent through the CABIN PRESSURE relay to CDC 1 and CDC 2. The CDCs remove power from the auto channels of the OFV. The outflow valve is driven closed using the manual channel of the IASC that was in control.ever
When the aircraft is on ground and the doors are open, the IASCs perform a MANUAL channel test as follows: •
The valve is moved from the fully open position to the fully closed position in 5 seconds
•
The valve is moved back to the fully open position in 5 seconds
•
If the valve position feedback does not match the commanded valve position, the MANUAL channel is reported as failed
A travel limiter device physically limits the outflow valve opening when extended. The travel limiter extends based on a preset cabin differential pressure, either in automatic or manual modes, when the aircraft flies over 17,000 ft. The travel limiter is powered by IASC 1. The travel limiter provides position feedback to IASC 1. The active IASC channel A software continuously monitors for any overpressure condition. If an overpressure condition is detected, the active IASC drives the OFV open. A force open command signal is generated for an emergency depressurization, or when the emergency ram air valve is opened. The force open command removes power from the auto channels, and drives the OFV open using the manual channel.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Cabin Pressure Relay Active
OFV Force Open Command
LEGEND CABIN PRESSURE RELAY
CDC EPC IASC OFV
Control and Distribution Cabinet Electrical Power Center Integrated Air System Controller Outflow Valve RS 422
Altitude Limit Auto 2 Manual Manual Close Manual Close Manual Open Manual Open Manual Position IASC 2
EPC 2 Cabin Pressure Relay Active Altitude Limit Manual Close Manual Open
CDC 2-6-12 Outflow VLV MAN DC ESS SSPC 3A BUS 2 Logic: Combinational Altitude Limit CDC 2-5-9 Outflow VLV AUTO 2 DC ESS SSPC 3A BUS 2
Manual Position
Logic: Combinational
CDC 2 Travel Limiter
Manual Position
Travel Limiter
Travel Limiter Position
Travel Limiter Position
FOV Force Open Command IASC 1
CDC 1-6-9 Outflow VLV AUTO 1 DC ESS SSPC 3A BUS 1 Logic: Combinational Altitude Limit
Auto 1 OFV
CDC 1
CS1_CS3_2130_015
Travel Limiter
Figure 102: Cabin Pressure Control System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
CABIN ALTITUDE LIMITATION ANNUNCIATION
CABIN ALTITUDE WARNING LOGIC
The baro-corrected landing elevation (L/EBARO) has three modes of operation for the cabin altitude limitation. One is for normal operation, and the other modes are for high altitude airport operations (HAAO).They are defined as follows:
The normal CAB ALT warning message threshold is 9865 ft.
•
Non-HAAO conditions defined for L/EBARO between -2000 ft and 8000 ft above sea level.
•
HAAO-1 conditions defined for L/EBARO between 8000 ft and 14,000 ft above sea level
•
HAAO-2 conditions are defined for L/EBARO between 14,000 ft and 14,500 ft above sea level
The cabin altitude caution and warning EICAS messages for non-HAAO operations, are set at: •
CAB ALT CAUTION 8500 ft
•
CAB ALT WARNING: 9865 ft
If the landing altitude is above 8000 ft: •
When aircraft starts the descent, the CABIN ALTITUDE warning starts to increase as a function of cab altitude caution threshold + 1365 ft and limited to 14,500 ft in HAAO-1 or 14,843 ft in HAAO-2.
If the takeoff is above 8,000 ft: •
When aircraft starts its climb and if aircraft altitude is above TOA + 5000 ft or 10 minutes after takeoff, the cabin altitude warning starts to decrease as a function of aircraft altitude to it nominal cruise set point of 9865 ft
In either landing or take off at high altitude, a CABIN ALT LEVEL HIGH advisory message is displayed to indicate that the cabin altitude warning threshold has been set above 9865 ft. The system limits cabin altitude to a maximum of 14843 ft in non-HAAO or HAAO-1 and 15200 ft in HAAO-2 in the event of a system malfunction.
CABIN ALTITUDE CAUTION LOGIC The normal CAB ALT CAUTION threshold is 8500 ft. If the landing altitude is above 8000 ft: •
When aircraft starts the descent and the aircraft altitude is below 37,000 ft, the CABIN ALTITUDE caution starts to increase as a function of aircraft altitude to a maximum of landing altitude + 500 ft and limited to 14,500 ft in HAAO-1 OR 14,843 ft in HAAO-2
The caution threshold increase rate is proportional to the A/C descent rate, aircraft altitude and L/EBARO. If the takeoff is above 8000 ft: •
When aircraft starts its climb and if aircraft altitude is above takeoff altitude (TOA) + 5000 ft or 10 minutes after takeoff, the CABIN ALTITUDE caution starts to decrease as a function of aircraft altitude to it nominal cruise set point of 8500 ft
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21 - Environmental Control 21-30 Cabin Pressure Control System
Caution Altitude
TOA Takeoff Altitude
CRUISE 41,000 ft
CRUISE 41,000 ft
Start of Cabin Climb
CAB ALT Warning TOA +1865 ft CAB ALT Caution TOA +500 ft 300 ft CAB ALT Warning 9865 ft
HAAO Landing Elevation
CAB ALT Warning TOA +1865 ft TOA CAB ALT Caution TOA +500 ft Takeoff Altitude 300 ft > 8000 ft
5000 ft
10 minutes
CAB ALT Caution 8500 ft
CAB ALT Warning 9865 ft CAB ALT Caution 8500 ft
Cabin Altitude Max 8000 ft
Cabin Altitude Max 8000 ft
Cruise
HAAO Takeoff Sequence
HAAO Descent
Climb
Cruise
Time
Time HAAO TAKEOFF CAB ALT LOGIC CS1_CS3_2130_012
HAAO LANDING CAB ALT LOGIC
Figure 103: Cabin Altitude Limitations Annunciation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
EMERGENCY DEPRESSURIZATION The cabin pressure can be reduced rapidly towards ambient pressure in an emergency. This function is known as emergency depressurization. Emergency depressurization is always available, whether the system is in AUTO mode, or in MANUAL mode. While in AUTO mode at the time of the emergency depressurization selection, a cabin rate of change limitation of 3500 ft/min is applied. In the MANUAL mode, the cabin rate of change limitation of 2500 ft/min is applied. The cabin altitude is limited to 14,500 ft. When the EMER DEPRESS PBA is pressed, the OFV is driven open at a rate that protects against excessive cabin rate of change. The ON legend on the PBA indicates the OFV is being driven open. An EMER DEPRESS ON caution message is displayed on EICAS. If the system was in automatic mode at the time of emergency depressurization selection, the function is accomplished by the active IASC channel A. If the system was in manual mode at the time of the emergency depressurization selection, a force open command is supplied through control and distribution cabinet 1 (CDC 1) to the safety channel. The outflow valve is driven open through the manual channel. The cabin altitude and rate limitations functions remain active. The CAB ALT caution and warning are also raised above the normal set point in order to monitor that the cabin altitude does not exceed 15,000 ft.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
DMC 1 and DMC 2
DITCHING AUTO PRESS FAIL
Channel A
Channel A
ON
MAN
MAN RATE
PAX OXY
DPLY DN
Channel B
UP
CDC 1
Channel B
CDC 2
PRESSURIZATION PANEL
Force Open Command Safety Channel
Force Open Command Safety Channel
CABIN PRESSURE RELAY
IASC 1
IASC 2
CDC DMC IASC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 RS 422
Auto 2
Manual OUTFLOW VALVE
Auto 1
CS1_CS3_2130_018
LEGEND
Figure 104: Emergency Depressurization Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
DITCHING The ditching mode is used to close all of the valves below the aircraft flotation line. Ditching can be selected in either AUTO or MANUAL pressurization modes. The DITCHING PBA closes all valves in preparation for a water landing. When the aircraft is below 15,000ft, pressing the guarded DITCHING PBA:
The following valves prevent water intake mechanically
•
Opens the outflow valve
•
Closes the flow control valves
The following valves are above the flotation line, but prevent water intake:
•
Trim air pressure-regulating valve
•
Safety valves
•
Trim air shutoff valve
•
Negative pressure-relief valve
•
Low-pressure ground connection
•
Bulkhead check valves
•
Water activated check valve
Closes all electrically controlled valves located below the flotation line •
Avionics exhaust valves
•
Ground valve
When pressing the DITCHING PBA, it latches and the ON legend illuminates white to indicate the electrically controlled valves have closed. If the ram air valve is open, the RAM AIR PBA is pressed to close it. A DITCHING MISCONFIG caution message indicates the DITCHING PBA is selected ON, and the ram air valve is open. When the cabin differential pressure has equalized, all valves are closed, and the airspeed is less than 80 kt, the outflow valve is driven closed. A DITCHING ON status message indicates the DITCHING PBA is pressed and the ditching sequence is complete.
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21 - Environmental Control 21-30 Cabin Pressure Control System
SFV (2X) – NPRV
AEV
GV
TAPRV TASOV
ERAV FCV
WACV LPGC
OFV
BCKV
AEV
ELECTRICALLY CLOSED VALVES
MECHANICALLY CLOSED VALVES
Forward AEV
BCKV
OFV
WACV
GV
LPGC
Mid AEV
SFV
FCV, TAPRV, TASOV
NPRV
AEV BCKV ERAV FCV GV LPGC NPRV SFV TAPRV TASOV WACV
Avionics Exhaust Valve Bulkhead Check Valve Emergency Ram Air Valve Flow Control Valve Ground Valve Low-Pressure Ground Connection Negative Pressure-Relief Valve Safety Valve Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Water Activated Check Valve
ERAV
CS1_CS3_2130_016
LEGEND
Figure 105: Ditching Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the cabin pressure control system:
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21 - Environmental Control 21-30 Cabin Pressure Control System
CAS MESSAGES
Table 32: ADVISORY Messages Table 30: WARNING Messages
MESSAGE
MESSAGE
LOGIC
CABIN DIFF PRESS
Positive differential pressure is above pneumatic relief setting.
CABIN ALT
Cabin altitude exceeds 10,000 ft under non-HAAO conditions, and up to 15,000 ft under HAAO conditions posted by IASC A channels. Cabin altitude exceeds 15,000 ft under any condition posted by IASC B channels. Safety cabin altitude high alarm from IASC1C or IASC2C.
Table 31: CAUTION Messages MESSAGE CABIN ALT
LOGIC No CABIN ALTITUDE warning message and cabin altitude exceeds 8,500 ft under non-HAAO conditions, and up to 15,000 ft under HAAO conditions. Cabin altitude exceeds 10,000 ft under HAAO conditions for more than 30 minutes.
AUTO PRESS FAIL
Both cabin pressure control AUTO functions are failed.
DITCHING MISCONFIG
DITCHING selected ON and RAM AIR selected OPEN.
LDG ELEV MISCONFIG
High airport altitude landing is selected and HAAO option is not present on aircraft.
EMER DEPRESS ON
EMER DEPRESS selected ON confirmed by either one auto mode or one manual mode.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
CABIN ALT LEVEL High airport altitude landing or takeoff sequence is HI initiated and the system operates in automatic mode. (increase of caution and warning threshold above normal setting). PRESSURIZATION FAULT
Loss of redundant or non-critical function for the pressurization system.
Table 33: STATUS Messages MESSAGE
LOGIC
DITCHING ON
DITCHING selected ON and sequence completed.
CABIN PRESS MAN
MANUAL PRESSURIZATION mode selected.
Table 34: INFO Messages MESSAGE
LOGIC
21 PRESSURIZATION FAULT - PRIM ALT LIM INOP
Loss of IASC 1 altitude limitation function.
21 PRESSURIZATION FAULT - BACKUP ALT LIM INOP
Loss of IASC 2 altitude limitation function.
21 PRESSURIZATION FAULT - CPCS AUTO MODE REDUND LOSS
One automatic mode failed.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Table 34: INFO Messages MESSAGE
LOGIC
21 PRESSURIZATION FAULT - MANUAL MODE INOP
OFV manual mode failed.
21 PRESSURIZATION FAULT - OFV FINGER FAIL OUT
Altitude limiter failed deployed.
21 PRESSURIZATION FAULT - OFV FINGER FAIL IN
Altitude limiter failed in closed position.
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ATA 28 - Fuel
BD-500-1A10 BD-500-1A11
28 - Fuel
Table of Contents 28-00 Fuel Safety ...................................................................28-2 Fuel safety History ............................................................28-2 25.981 Fuel Tank Ignition Prevention ..........................28-2 Special Federal Aviation Regulation 88 .......................28-2 European Aviation Safety Agency Regulations ...........28-2 Airworthiness Limitations-Fuel System Limitations ..........28-4 Critical Design Configuration Control Limitations .............28-6 CDCCL Warning ...............................................................28-8 Fuel Tank Access and Venting .......................................28-10 Fuel Tank Access ......................................................28-10 Fuel Tank Venting......................................................28-10 Fuel Tank Entry ..............................................................28-12 Preparation ................................................................28-12 Entry...........................................................................28-12 28-11 Fuel Storage ...............................................................28-14 General Description ........................................................28-14 Component Location .....................................................28-16 Flapper Check Valves................................................28-16 Fuel Drain Valves.......................................................28-16 Fuel Tank Access Panels ..........................................28-16 Practical Aspects ............................................................28-18 Fuel Drain Valve Servicing.........................................28-18 28-12 Vent System ...............................................................28-20 General Description .......................................................28-20 Component Location ......................................................28-22 Float Drain Valves......................................................28-22 Center Tank Float Vent Valves ..................................28-22 Wing Tank Float Vent Valves.....................................28-24 Bifurcated Vent Openings ..........................................28-24 NACA Inlet Scoop ......................................................28-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Vent Relief Flame Arrestor ........................................ 28-26 Positive Pressure-Relief Valves ................................ 28-28 Center Tank Positive Pressure-Relief Valve ............. 28-28 Center Tank Negative Pressure-Relief Valve ........... 28-28 28-23 Refuel/Defuel System ................................................ 28-30 General Description ....................................................... 28-30 Refueling ................................................................... 28-32 Defueling ................................................................... 28-36 Component Location .................................................... 28-40 Refuel Adapter .......................................................... 28-40 Refuel/Defuel Control Panel...................................... 28-40 Refuel Shutoff Valve ................................................. 28-42 Refuel Control Solenoid Valve .................................. 28-42 Defuel/Isolation Transfer Shutoff Valve and Actuator ....................................... 28-44 Manual Defuel Shutoff Valve and Lever.................... 28-44 Component Information ................................................. 28-46 Refuel Adapter .......................................................... 28-46 Refuel/Defuel Control Panel...................................... 28-48 Controls and Indications ............................................... 28-50 Refuel/Defuel Control Panel...................................... 28-50 Optional Virtual Refuel Panel .................................... 28-54 Operation ....................................................................... 28-56 Auto Refuel Operation............................................... 28-56 Manual Refuel Operation .......................................... 28-58 Pressure Defuel Operation........................................ 28-60 Suction Defuel Operation .......................................... 28-62 Detailed Description ..................................................... 28-64 Refuel Power............................................................. 28-64 Practical Aspects ........................................................... 28-66 Onboard Maintenance System.................................. 28-66
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28 - Fuel
28-21 Engine Feed System ..................................................28-70 General Description .......................................................28-70 Component Location .....................................................28-72 Engine Feed Ejector Pump ........................................28-72 AC Boost Pump .........................................................28-72 Engine Feed Shutoff Valve and Actuator...................28-72 Engine Feed Pressure Switch....................................28-72 Component Information .................................................28-74 AC Boost Pump .........................................................28-74 Controls and Indications ................................................28-76 Engine Fire PBA ........................................................28-76 Fuel Panel..................................................................28-76 Synoptic Display ........................................................28-78 Detailed Description ......................................................28-80 Left Engine Feed........................................................28-80 Right Engine Feed .....................................................28-82 Monitoring and Tests .....................................................28-84 CAS Messages ..........................................................28-85 28-21 APU Feed System ......................................................28-86 General Description .......................................................28-86 Component Location ......................................................28-88 APU Feed Shutoff Valve and Actuator.......................28-88 APU Feed Line, Shroud, and Drain Masts.................28-90 Controls and Indications .................................................28-92 APU Start Switch .......................................................28-92 APU Fire PBA ............................................................28-92 Detailed Description ......................................................28-94 Monitoring and Tests .....................................................28-96 CAS Messages ..........................................................28-97
Automatic Transfer ................................................. 28-102 Center Tank to Main Tank Fuel Transfer ................ 28-104 Manual Fuel Transfer .............................................. 28-106 Component Location ................................................... 28-112 Scavenge Ejector Pump.......................................... 28-112 Transfer Float Valve................................................ 28-114 Transfer Ejector Pump ............................................ 28-114 Gravity Crossflow Shutoff Valve.............................. 28-116 Controls and Indications ............................................. 28-118 Synoptic Page ......................................................... 28-120 Monitoring and Tests .................................................. 28-122 CAS Messages ....................................................... 28-123 28-40 Fuel Indicating System ............................................ 28-124 General Description .................................................... 28-124 Component Location .................................................. 28-126 Fuel Probes............................................................. 28-126 Temperature Sensors ............................................. 28-126 Remote Data Concentrator ..................................... 28-128 Fuel Quantity Computer .......................................... 28-130 Fuel Probes............................................................. 28-132 Temperature Sensor ............................................... 28-132 Detailed Component Information ................................ 28-134 Fuel Quantity Computer Interface ........................... 28-134 Controls and Indications ............................................. 28-136 Detailed Description ................................................... 28-138 Monitoring and Tests .................................................. 28-140 CAS Messages ....................................................... 28-141 Practical Aspects ........................................................ 28-144 Onboard Maintenance System ............................... 28-144
28-22 Fuel Transfer System .................................................28-98 General Description .......................................................28-98 Scavenge System ....................................................28-100 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel
List of Figures Figure 1: Fuel Safety History ................................................28-3 Figure 2: Airworthiness Limitations-Fuel System Limitations ............................................................28-5 Figure 3: Critical Design Configuration Control Limitation ..................................................28-7 Figure 4: Critical Design Configuration Control Limitation Warning in Aircraft Maintenance Publication ..........................28-9 Figure 5: Fuel Tank Access and Venting............................28-11 Figure 6: Fuel Tank Entry ...................................................28-13 Figure 7: Fuel Storage System...........................................28-15 Figure 8: Fuel Storage System Component Location.........28-17 Figure 9: Fuel Drain Valve Servicing ..................................28-19 Figure 10: Fuel Vent System ................................................28-21 Figure 11: Float Drain Valve and Center Tank Float Vent Valve ..................................................28-23 Figure 12: Bifurcated Vent Openings and Wing Tank Float Vent Valve................................28-25 Figure 13: NACA Inlet Scoop and Vent Relief Flame Arrestor.....................................................28-27 Figure 14: Positive Pressure-Relief Valves, Center Tank Positive, and Center Tank Negative Pressure-Relief Valves.........................28-29 Figure 15: Refuel/Defuel System..........................................28-31 Figure 16: Auto Refueling.....................................................28-33 Figure 17: Manual Refueling ................................................28-35 Figure 18: Suction Defueling ................................................28-37 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 19: Pressure Defueling ............................................. 28-39 Figure 20: Refuel Adapter and Refuel/Defuel Control Panel Location................. 28-41 Figure 21: Refuel Control Solenoid Valve and Refuel Shutoff Valve Component Location......... 28-43 Figure 22: Defuel System Components .............................. 28-45 Figure 23: Refuel Adapter.................................................... 28-47 Figure 24: Refuel/Defuel Control Panel ............................... 28-49 Figure 25: Refuel/Defuel Control Panel ............................... 28-51 Figure 26: Refuel/Defuel Control Panel Messages.............. 28-53 Figure 27: Virtual Refuel Panel ............................................ 28-55 Figure 28: Automatic Refuel Operation................................ 28-57 Figure 29: Manual Refuel Operation.................................... 28-59 Figure 30: Pressure Defuel Operation ................................. 28-61 Figure 31: Suction Defuel Operation.................................... 28-63 Figure 32: Refuel Power (Normal Operation - Door Closed)...................... 28-65 Figure 33: Refuel Solenoid Command Test......................... 28-67 Figure 34: Defuel Shutoff Valve Command Test ................. 28-69 Figure 35: Engine Feed System .......................................... 28-71 Figure 36: Engine Feed Component Location ..................... 28-73 Figure 37: AC Boost Pump .................................................. 28-75 Figure 38: Engine Fire PBAs and Engine Fuel Panel .......... 28-77 Figure 39: Fuel Synoptic Page............................................. 28-79 Figure 40: Left Engine Feed Detailed Description ............... 28-81 Figure 41: Right Engine Feed Detailed Description............. 28-83
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Figure 42: APU Fuel Feed System.......................................28-87 Figure 43: APU Feed Shutoff Valve and Actuator ................28-89 Figure 44: APU Fuel Line, Shroud and Drain Masts ............28-91 Figure 45: APU Start Switch and Fire PBA ..........................28-93 Figure 46: APU Feed Shutoff Valve Detailed Description ............................................28-95 Figure 47: Fuel Transfer System ..........................................28-99 Figure 48: Scavenge System .............................................28-101 Figure 49: Automatic Fuel Transfer ....................................28-103 Figure 50: Center Tank to Main Tank Fuel Transfer ..........28-105 Figure 51: Main Tank to Main Tank Fuel Transfer .............28-107 Figure 52: Main Tank to Center Tank Fuel Transfer ..........28-109 Figure 53: Gravity Fuel Transfer.........................................28-111 Figure 54: Scavenge Ejector Pump....................................28-113 Figure 55: Center Tank to Main Tank Fuel Transfer Ejector Pump and Float Valve ............28-115 Figure 56: Gravity Crossflow Shutoff Valve........................28-117 Figure 57: Fuel Transfer Controls.......................................28-119 Figure 58: Fuel Synoptic Page ...........................................28-121 Figure 59: Fuel Quantity Indication.....................................28-125 Figure 60: Fuel Quantity Probes and Temperature Sensors........................................28-127 Figure 61: Remote Data Concentrator ...............................28-129 Figure 62: Fuel Quantity Computer ....................................28-131 Figure 63: Fuel Probes and Temperature Sensor ..............28-133 Figure 64: Fuel Quantity Interface ......................................28-135 Figure 65: Fuel Controls and Indications............................28-137 Figure 66: Fuel Quantity System ........................................28-139 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 67: Low-Level Discrete Output Test ....................... 28-145 Figure 68: Clear NVM Function ......................................... 28-147
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FUEL - CHAPTER BREAKDOWN Fuel Chapter Breakdown
Fuel Safety
Engine Feed System
1
Fuel Storage
5
APU Feed System
2
Vent System
6
Fuel Transfer System
3
Pressure Refuel/Defuel System
4
7
Fuel Indicating System
8
28 - Fuel
28-00 FUEL SAFETY FUEL SAFETY HISTORY Since 1959 there have been 17 fuel tank ignition events, resulting in 542 fatalities, 11 hull losses, and 3 others with substantial damage to the aircraft. Since the 1960's, there have been 5 key accidents involving fuel tank explosions which call into question the fundamental safety strategy applied to fuel systems of large commercial airplanes. Because of this accident, multiple airworthiness directives (ADs) on several aircraft types were issued. 25.981 FUEL TANK IGNITION PREVENTION This advisory circular (AC) 25.981 fuel tank ignition prevention provides guidance for demonstrating compliance with the certification requirements for prevention of ignition sources within the fuel tanks of transport category airplanes. The TWA 800 accident brought the realization that some fuel tanks could be flammable for much of their operational time. In their efforts to determine the cause of TWA 800 accident, the FAA concluded that many current airplanes had problems in their ignition prevention approaches. An additional independent layer of protection would be needed to back up the ignition prevention strategy. SPECIAL FEDERAL AVIATION REGULATION 88
A flammability reduction program would be necessary to prevent the operation of transport category airplanes with explosive fuel-air mixtures in the fuel tank. These fuel tanks were called high flammability tanks. The overall goal was to make sure flammability reduction of high flammability tanks became an integral part of aircraft safety. EUROPEAN AVIATION SAFETY AGENCY REGULATIONS A similar regulation had been recommended by the European Joint Aviation Authorities (JAA) to the European national aviation authorities in JAA letter 04/00/02/07/03-L024 of February 3, 2003. The review was mandated by European national airworthiness authorities using JAR 25.901(c), and 25.1309. In August 2005, the European Aviation Safety Agency (EASA) published a policy statement on the process for developing instructions for maintenance and inspection of fuel tank system ignition source prevention that also included the EASA expectations with regard to compliance times of the corrective actions. The EASA policy statement reset the date for compliance of unsafe items to July 1, 2006. Fuel airworthiness limitations are items from a systems safety analysis that have been shown to have failure modes associated with an unsafe condition as defined in SFAR 88. These are identified in failure conditions for which an unacceptable probability of ignition risk could exist if specific tasks and/or practices are not performed in accordance with the manufacturers requirements.
In response to these findings, the FAA issued Special Federal Aviation Regulation (SFAR) No. 88 in June of 2001. SFAR 88 required a safety review of the aircraft fuel tank system to determine that the design meets the requirements of FAR 25.901 and 25.981 (a) and (b).
This EASA airworthiness directive mandated the fuel airworthiness limitation critical design configuration control limitations (CDCCLs) consisting of maintenance and inspection tasks for the type of aircraft, that resulted from the design reviews, JAA recommendations, and the EASA policy statement.
The safety reviews were very valuable in that they revealed unexpected ignition sources. There was difficulty in identifying all ignition sources and a number of previously unknown failures were found.
The FAA and EASA have also mandated that operators must train their maintenance and engineering personnel regarding the changes brought about by SFAR 88.
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28 - Fuel
Aircraft
Location
Year
Cause
Boeing 707
Elkton, Maryland
1963
Lightning
Boeing 747
Madrid, Spain
1976
Lightning
Boeing 737
Manila, Philippines
1990
Unconfirmed
Boeing 747
New York, New York
1996
Center fuel tank explosion. Exact cause undetermined.
Boeing 737
Bangkok, Thailand
2001
Center fuel tank explosion. Possible center fuel pump failure.
TWA 800 B747
CS1_CS3_2800_010
MAJOR ACCIDENTS RELATED TO FUEL TANK EXPLOSIONS
Figure 1: Fuel Safety History Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-00 Fuel Safety
AIRWORTHINESS LIMITATIONS-FUEL SYSTEM LIMITATIONS Airworthiness limitation items (ALI) are mandatory maintenance items for the fuel system that can include CDCCLs, inspections, or other procedures necessary to make sure the fuel tank flammability exposure does not increase above the certification limits as a result of maintenance actions, repairs, or alterations throughout the operational life of the airplane. The Airworthiness Limitations (AWL) gives the mandatory scheduled maintenance requirements applicable to the C Series aircraft. The fuel system limitation (FSL) information is found in the AWL section 7. This section contains task data about the aircraft fuel system maintenance tasks generated by the fuel tank system safety and ignition prevention analysis accomplished in compliance with federal aviation administration (FAA) Special Federal Aviation Regulation (SFAR) No. 88 - fuel tank system fault tolerance evaluation requirements. FSL tasks specified are mandatory airworthiness limitations and must be incorporated into the operators locally approved maintenance schedule.
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CS1_CS3_2800_005
28-00 Fuel Safety
Figure 2: Airworthiness Limitations-Fuel System Limitations Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-00 Fuel Safety
CRITICAL DESIGN CONFIGURATION CONTROL LIMITATIONS Critical design configuration control limitations (CDCCLs), a type of airworthiness limitation, include those features of the design that must be present and maintained to achieve the safety level intended by FAR 25.981 for the operational life of the airplane. The information is found in the airworthiness limitations (AWL) section 7. This section contains information about the critical design configuration control limitations (CDCCLs) generated by the fuel tank system safety and ignition prevention analysis accomplished in compliance with FAA SFAR 88 - fuel tank system fault tolerance evaluation requirements. The critical design configuration control limitations (CDCCLs) are established to preserve the critical ignition source prevention features of the fuel system design that are necessary to prevent the occurrence of an unsafe condition identified by the fuel tank safety assessment (FTSA) analysis. The purpose of the CDCCLs is to provide instructions to aircraft maintenance personnel to retain the critical ignition source prevention features during the time of configuration change that may be caused by alterations, repairs or maintenance actions. Identification of the CDCCLs design features, parts or components is provided within the applicable task procedures of the instructions for continued airworthiness (ICA), including the Aircraft Maintenance Publication (AMP). The critical ignition source prevention features of the parts or components identified in the list of CDCCLs shown in Table 2 must be maintained in a configuration identical to the approved type design for the C Series aircraft.
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CS1_CS3_2800_006
28-00 Fuel Safety
Figure 3: Critical Design Configuration Control Limitation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-00 Fuel Safety
CDCCL WARNING Identification of the CDCCL design features, parts, or components is provided within the applicable task procedures of the instructions for continued airworthiness (ICA), including the Aircraft maintenance Publication (AMP). The following warning is included when a task can impact CDCCL features.
WARNING DO NOT DAMAGE OR MODIFY THE FUEL TANK PLUMBING LINES, SELF-BONDING COUPLINGS, CONDUCTIVE FITTINGS, AND METAL-TO-METAL INTERFACE ELECTRICAL BONDING OF THE FUEL TANK COMPONENTS. MAKE SURE THAT PROTECTIVE SEALANT IS RE-APPLIED AND/OR BONDING CHECK IS DONE AS INSTRUCTED. THESE DESIGN FEATURES ARE CLASSIFIED AS CRITICAL DESIGN CONFIGURATION CONTROL LIMITATION (CDCCL) ITEMS. THEY GIVE ELECTROSTATIC AND LIGHTNING PROTECTION. DAMAGE OR MODIFICATION TO THESE ITEMS CAN POSSIBLY CAUSE AN IGNITION SOURCE INTO THE FUEL TANK BECAUSE OF ELECTROSTATICS OR A LIGHTNING STRIKE.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-8
CS130_28_L2_FuelSafety_CCDLWarning.fm
28 - Fuel
CS1_CS3_2800_007
28-00 Fuel Safety
Figure 4: Critical Design Configuration Control Limitation Warning in Aircraft Maintenance Publication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-9
CS130_28_L2_FuelSafety_CCDLWarning.fm
28 - Fuel 28-00 Fuel Safety
FUEL TANK ACCESS AND VENTING FUEL TANK ACCESS
The following is a summary of steps to remove the fuel fumes from the fuel tanks using the fuel tank venting kit:
The following is a summary of steps to prepare the fuel tank for access:
1. Put an air compressor more than 30 m (100 ft) from the aircraft.
1. Disconnect the aircraft batteries.
2. Connect the air compressor to the same ground point to which the aircraft is connected.
2. Disconnect external power and put warning placards in the flight deck ‘Do not apply electrical power to the aircraft’. 3. Establish a safety zone around the aircraft and install warning signs ‘Open fuel tank around the aircraft’. 4. After the tanks are defueled, drain the remaining fuel through the water drain valves. 5. Remove the fuel fumes from the fuel tank using the fuel tank venting kit.
3. Install the inlet wing adapter at the outboard access panel opening. 4. Connect air hose from the air compressor to the inlet wing adapter. 5. Install the outlet wing adapter at the inboard access panel opening. 6. Connect the exhaust hose to the outlet wing adapter and to a ground point. Put the exhaust hose outlet so it discharges outside of the hangar.
6. After 30 minutes, monitor the fuel tank atmosphere using a gas detector. 7. Do not enter the fuel tank until the fuel fume concentration is less than the lower explosion limit and the oxygen level is at least 19.5%. 8. Continue to vent the fuel tank when work is being done inside the fuel tank. 9. When working in the fuel tank, measure the fuel fume concentration in and around the fuel tank every 30 minutes. 10. Only personnel trained to enter a fuel tank are allowed to enter the fuel tank. 11. Remove any residual fuel using cotton cloths or sponges.
WARNING 1. MAKE SURE NO PERSON ENTERS THE FUEL TANK IF THE FUME CONCENTRATION IS MORE THAN THE LOWER EXPLOSION LIMIT. 2. MAKE SURE THAT ALL EQUIPMENT IS CLEAR OF THE EXHAUST DUCT. STATIC ELECTRICITY GENERATED BY AIRFLOW FROM THE EXHAUST DUCT CAN CAUSE AN EXPLOSION. 3. OBEY ALL LOCAL REGULATIONS WHEN DISCARDING FUEL.
FUEL TANK VENTING The fuel tank venting kit has the necessary equipment to ventilate the fuel tank including air hoses, air movers, antistatic ducts, couplings, and wing adapters. Complete instructions are available with the kit.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-10
CS130_28_L2_FuelSafety_FuelTankAccess.fm
28 - Fuel 28-00 Fuel Safety
Wing Adaptors
Wing Venting Kit
Couplings
100 mm Antistatic Ducts
Small Air Mover
200 mm to 100 mm Coupling 200 mm Antistatic Ducts
CS1_CS3_2800_008
Large Air Mover
Air Hoses
Figure 5: Fuel Tank Access and Venting Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-11
CS130_28_L2_FuelSafety_FuelTankAccess.fm
28 - Fuel 28-00 Fuel Safety
FUEL TANK ENTRY The following is a summary of steps to enter the fuel tank:
PREPARATION The following is a summary of steps to prepare for fuel tank entry: 1. Obey all local safety regulations during fuel system maintenance. 2. Firefighting extinguishers must be positioned near the aircraft. 3. Connect the grounding cable to an approved ground and to an aircraft ground point. 4. Ground the maintenance stands and equipment to the same ground point the aircraft is connected to. 5. Do not connect or disconnect electrical equipment to or from energized outlets that are less than 30.5 m (100 ft) from an open fuel tank . 6. Use explosion-proof floodlights and extension cords for external lighting. Use explosion-proof flashlights or explosion-proof lights inside the fuel tank. Explosion-proof lights must be connected to the electrical power before being put it in the fuel tank.
NOTE Only personnel with fuel tank entry training should enter a fuel tank. The local authorities may require the personnel to have a tank entry permit
1. Wear clean lint-free cotton coveralls with buttons or zippers that do not cause sparks in the fuel tanks. Do not use wool, silk, nylon, or other synthetic clothing. When entering the fuel tank, use cotton hair and shoe covers as well. 2. When entering a fuel tank, an observer must be stationed near the fuel tank opening. The observer must remain in contact with personnel inside the tank and monitor their condition. 3. Metal boxes or containers with sharp or rough edges can cause damage to the fuel tank. Use containers with rounded corners that do not generate static electricity. Sharp-edged tools must be put it in a container when not in use. 4. Use only sponges or cotton rags to absorb fuel. Other materials can cause a static discharge 5. Leave the fuel tank immediately if the fuel concentration exceeds the lower explosion limit or is below the minimum oxygen concentration level. 6. Keep a list of all the tools, equipment, and materials entering the fuel tank. All tools, equipment, and material must be accounted for before the tank is closed. 7. Make sure the fuel tank is thoroughly inspected for cleanliness before closing the fuel tank.
ENTRY Before entering a fuel tank, check that the breathing apparatus kit is in good working condition. The breathing apparatus kit has an instruction manual, pressure regulator and valve, air filter, air hoses, and the necessary accessories to supply air to a full or half face mask.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
WARNING DO NOT USE STEEL WOOL OF ANY TYPE IN THE FUEL TANKS. PIECES OF STEEL WOOL CAN BREAK OFF AND CAUSE DAMAGE OR AN EXPLOSION IF THEY TOUCH ELECTRICAL COMPONENTS.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-12
CS130_28_L2_FuelSafety_FuelTankEntry.fm
28 - Fuel 28-00 Fuel Safety
Breathing Tube Panel Filter and Regulator
Regulator Valve Air Hose
FIRE EXTINGUISHER
COTTON COVERALLS
Face Mask
EXPLOSION PROOF LIGHTS
BREATHING APPARATUS KIT COTTON RAGS
CS1_CS3_2800_009
Half Mask
Figure 6: Fuel Tank Entry Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-13
CS130_28_L2_FuelSafety_FuelTankEntry.fm
28 - Fuel 28-11 Fuel Storage
28-11 FUEL STORAGE GENERAL DESCRIPTION The aircraft fuel is stored in the integral center, right wing, and left wing tanks located within the wing box structure. The tanks are isolated by a common boundary formed by structural ribs, and have independent provisions for fuel gauging and refueling. Baffle ribs are installed in the wing to restrict fuel migration outboard during aircraft maneuvers. The wings and the center wing box are all constructed using carbon fiber reinforced polymer (CFRP), with the exception of the wing ribs, which are metallic.
One flapper check valve allows fuel to drain from the surge tank back to the main tank. Two fuel drain valves are installed in each of the fuel tanks at low points to provide water drainage.
Each wing tank has a main and a collector tank. The collector tank is partially sealed and is located inboard of the wing tank. A dry bay is located above each engine strut in the wing tanks. The dry bays prevent fuel from spilling onto an engine if a wing tank is ruptured. Access panels, located on the underside of the wing, provide access for the inspection, repair, and replacement of components inside the wing tanks. All ribs, including the baffle ribs, incorporate drain passages to prevent the accumulation of water and trapped fuel. Each rib has openings at the upper intersection with the wing inner-skin liner, which facilitates airflow within the tank. The interior of the tanks is chemically treated against corrosion and coated with a biocide compound. All joints and rivet areas are sealed to ensure tank seal integrity. Flapper check valves allow fuel to gravity feed from the main tanks to the collector tank to ensure that fuel is available at the fuel pumps in all aircraft attitudes. They also prevent the flow of fuel from the collector tank to the main tank during wing down and uncoordinated aircraft maneuvers. The check valve is a rubberized fabric hinged flapper.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-14
CS130_28_L2_FuelStorage_General.fm
28 - Fuel 28-11 Fuel Storage
RIB 5
RIB 0
Fuel Tank Access Panel (typical)
RIB 21 RIB 7
Fuel Tank Access Panel (typical)
RIB 21 RIB 23
RIB 7
Fuel Drain Valve
Flapper Check Valve
Surge Tank Main Tank Wing Dry Bay Collector Tank Center Tank
Flapper Check Valves (4x)
NOTE Fuel Drain Valve
Components from right wing shown. Left wing similar.
CS1_CS3_2811_007
LEGEND
Figure 7: Fuel Storage System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-15
CS130_28_L2_FuelStorage_General.fm
28 - Fuel 28-11 Fuel Storage
COMPONENT LOCATION The following components are installed in the fuel storage system: •
Flapper check valves
•
Fuel drain valves
•
Fuel tank access panels
FLAPPER CHECK VALVES Four flapper check valves are located on RIB 7, between the main and collector tanks on each wing. One flapper check valve is located on RIB 21, between the main tank and surge tank on each wing. FUEL DRAIN VALVES Fuel drain valves are installed at low points in the center, main, and collector tanks. The fuel drain valves are installed between the RIB 1 and RIB 2 of the left and right side of the center wing box. There is one fuel drain valve between RIB 5 and RIB 6, and another between RIB 7 and RIB 8 on each wing. FUEL TANK ACCESS PANELS Fuel tank access panels are installed on the underside of the wings and center tank, between each rib from RIB 0 to RIB 23.
WARNING BE CAREFUL WHEN YOU REMOVE THE ACCESS PANEL. FUEL CAN BE ON THE INNER SURFACE OF THE PANEL. FUEL CAN CAUSE INJURY TO PERSONS AND/OR DAMAGE TO EQUIPMENT.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-16
CS130_28_L2_FuelStorage_CompLocation.fm
28 - Fuel 28-11 Fuel Storage
RIB 5 RIB 0
Fuel Tank Access Panel (typical)
Fuel Tank RIB 7 Access Panel (typical) RIB 21 RIB 23
FUEL TANK ACCESS PANELS
Surge Tank Main Tank Wing Dry Bay Collector Tank Center Tank
FLAPPER CHECK VALVE
FUEL DRAIN VALVE
CS1_CS3_2811_002
LEGEND
Figure 8: Fuel Storage System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-17
CS130_28_L2_FuelStorage_CompLocation.fm
28 - Fuel 28-11 Fuel Storage
PRACTICAL ASPECTS FUEL DRAIN VALVE SERVICING The fuel drain valves are used for: •
Removing water and contaminants
•
Fuel sampling
•
Draining the remaining fuel after defueling a tank
The fuel drain valve is a simple poppet operated type device. O-ring seals on the valves may be replaced without removing the valve from the tank. The valve has a push-to-drain operation and can be locked in the open position using a standard screwdriver.
NOTE The fuel drain valve body is plastic. Use caution when attempting to operate in cold weather.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-18
CS130_28_L2_FuelStorage_PracticalAspects.fm
28 - Fuel 28-11 Fuel Storage
SERVICE POSITION FOR O-RING REPLACEMENT CS1_CS3_2811_006
OPEN FOR DRAINING
Figure 9: Fuel Drain Valve Servicing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-19
CS130_28_L2_FuelStorage_PracticalAspects.fm
28 - Fuel 28-12 Vent System
28-12 VENT SYSTEM GENERAL DESCRIPTION The fuel vent system maintains the pressure of the fuel tanks near the pressure of the outside atmosphere to prevent damage to the wing structure. The surge tanks are open to outside air through the vent scoops. During flight, the vent scoops maintain positive pressure inside the surge tanks. Each fuel tank is connected to a wing surge tank through a vent tube. The vent tubes have bifurcated openings at the two ends to provide added protection against blockage. The right wing tank vents through the right surge tank, and the left wing tank and the center tank vent through the left surge tank. The vent tubes keep the pressure of all the fuel tanks near the pressure in the surge tanks. Float vent valves and drains ensure the vent tubes are not blocked by fuel that might enter the lines. They also drain any fuel that may have accumulated in the vent tubes. The float drain valves close when the fuel level reaches the vent tubes. The center tank float drain valves also provide additional venting redundancy for the center tank. Center tank float vent valves are installed in the center tank to ensure venting in case the vent tube openings are submerged in fuel during aircraft maneuvers. The main tank float vent valve remains above the fuel level in any ground condition due to the aircraft wing dihedral. Any fuel entering the surge tank through the float vent valve drains back to the main tank through the surge tank flapper check valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The flow of air between the surge tanks and the atmosphere is normally through the NACA inlet scoop and flame arrestor. The inlet scoop creates a small pressure within the surge tank during flight. The inlet scoop also provides a path for fuel to drain overboard should the surge tank overflow with fuel. The flame arrestor ensures that a flame does not enter the fuel tanks through the vent system. In case of a blockage, the flame arrestor has a built-in negative pressure-relief valve to provide backup pressure equalization. In normal operation, air flows between the main wing tanks and surge tanks through the bifurcated vent openings or through the float vent valves, depending on the aircraft attitude and the level of fuel in the tank. For the center tank vent, air flows through the vent tubes, either by the bifurcated opening or through one of the float vent valves installed in the vent tube. Main wing tank pressure-relief valves allow pressure relief to the atmosphere. Center tank positive pressure is relieved into the right wing tank and negative pressure is relieved into the right surge tank. Surge tank flapper valves also provide redundant negative pressure relief to the main tanks. If the flow path is submerged by fuel and the alternate path is blocked, the internal pressure of the tank will cause fuel to be spilled into the surge tank. The fuel either drains back through the flapper valves or spills into the atmosphere through the NACA inlet scoop and flame arrestor. In case of a complete blockage in the normal flow path, the appropriate pressure-relief valve (PRV) will provide backup pressure equalization.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-20
CS130_28_L2_VentSystem_General.fm
28 - Fuel 28-12 Vent System
Center Tank Vent Tube
Left Main Tank Vent Tube
Center Tank Positive Pressure-Relief Tube
L SURGE TANK
CENTER TANK
L MAIN TANK
Right MainTank Vent Tube
R DRY BAY
L DRY BAY
Bifurcated Opening
Center Tank Negative Pressure-Relief Tube
COLLECTOR TANK
R MAIN TANK COLLECTOR TANK
R SURGE TANK
Negative Pressure-Relief Valve
LEGEND
CS1_CS3_2812_002
NACA Inlet Scoop Flame Arrestor Float Vent Valve Pressure-Relief Valve Float Drain Valve Flapper Check Valve
Figure 10: Fuel Vent System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-21
CS130_28_L2_VentSystem_General.fm
28 - Fuel 28-12 Vent System
COMPONENT LOCATION The following components are installed in the fuel vent system: •
Float drain valves
•
Center tank float vent valves
•
Wing tank float vent valves (refer to figure 12)
•
Bifurcated vent openings (refer to figure 12)
•
NACA inlet scoop (refer to figure 13)
•
Vent relief flame arrestor (refer to figure 13)
•
Positive pressure-relief valves (refer to figure 14)
•
Center tank negative pressure-relief valve (refer to figure 14)
FLOAT DRAIN VALVES Float drain valves are installed along the vent tubes in the center and wing tanks at RIB 2 and RIB 9. CENTER TANK FLOAT VENT VALVES Two float vent valves are installed in the center tank between RIB 0 and RIB 1 and at RIB 6 of the left wing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-22
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 0 RIB 6 RIB 9
FLOAT DRAIN VALVE
CENTER TANK FLOAT VENT VALVE
CS1_CS3_2812_006
RIB 9
Figure 11: Float Drain Valve and Center Tank Float Vent Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-23
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
WING TANK FLOAT VENT VALVES Wing tank float vent valves, installed near the top of RIB 21, vent each main tank into the surge tank. BIFURCATED VENT OPENINGS The bifurcated vent openings are located in the surge tanks at RIB 22, RIB 8, and at RIB 6R in the right wing only.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-24
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 6 RIB 8 RIB 8
BIFURCATED VENT OPENINGS
WING TANK FLOAT VENT VALVE
CS1_CS3_2812_007
RIB 22
RIB 22
Figure 12: Bifurcated Vent Openings and Wing Tank Float Vent Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-25
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
NACA INLET SCOOP A NACA inlet scoop is installed in each surge tank outboard of RIB 22. VENT RELIEF FLAME ARRESTOR The vent relief flame arrestor is installed in the left and the right wings between RIB 22 and RIB 23.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-26
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 22 RIB 23
Flame Arrester
NACA VENT SCOOP AND FLAME ARRESTOR
CS1_CS3_2812_009
Negative Pressure-Relief Valve
Figure 13: NACA Inlet Scoop and Vent Relief Flame Arrestor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-27
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
POSITIVE PRESSURE-RELIEF VALVES Positive pressure-relief valves are installed at the access panel on the left and the right wings between RIB 20 and RIB 21. CENTER TANK POSITIVE PRESSURE-RELIEF VALVE The positive pressure-relief valve for the center tank is installed on RIB 7 in the right wing tank. CENTER TANK NEGATIVE PRESSURE-RELIEF VALVE The center tank negative pressure-relief valve is installed on the right wing at RIB 21.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-28
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 7
RIB 21
RIB 21
CENTER TANK POSITIVE PRESSURE-RELIEF VALVE
CENTER TANK NEGATIVE PRESSURE-RELIEF VALVE
CS1_CS3_2812_008
POSITIVE PRESSURE-RELIEF VALVE
Figure 14: Positive Pressure-Relief Valves, Center Tank Positive, and Center Tank Negative Pressure-Relief Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-29
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
28-23 REFUEL/DEFUEL SYSTEM GENERAL DESCRIPTION Aircraft refueling and defueling is carried out using a single-point refuel/defuel adapter, located at the right wing. The refuel adapter connects to the refuel manifold that runs near the rear spar of the fuel tanks. The refuel/defuel tube connects the refuel adapter to all of the tanks via the refuel shutoff valves.
The refuel adapter must be set to the defuel position before connecting the fuel truck hose. Defueling the aircraft can only be done in MANUAL mode.
Each tank has one refuel shutoff valve and they connect the refuel/defuel tube to the tanks. The refuel control solenoid valve controls the opening and closing of the refuel shutoff valve. The refuel control solenoid valve is controlled by the fuel quantity computer (FQC). When the solenoid is energized, fuel pressure is directed to open the refuel shutoff valve. When the refuel shutoff valve opens, fuel discharges into the tank through a diffuser. Refueling and defueling is controlled from the refuel/defuel control panel (RDCP) located on the fuselage at right wing root. Pressure refueling of the aircraft is carried out in manual or automatic mode, through the appropriate AUTO/MANUAL switch selection on the RDCP. The recommended fuel pressure for refueling is 50 psig. The FQC receives power from DC ESS BUS 1 and DC ESS BUS 2 during normal operation. The refuel shutoff valves receive power from DC ESS BUS 1 during normal operation. When the RDCP door panel is opened, the power for the FQC and the refuel shutoff valves is switched to the DC EMER BUS. The RDCP is powered by the FQC through the door switch. A second contact on the door switch provides a signal that the RDCP door is open. The aircraft is defueled by connecting the refueling manifold to the engine feed manifold. The defuel/isolation transfer shutoff valve connects the manifolds. The valve is powered and controlled by the FQC. When DEFUEL is selected on the RDCP, the valve opens and DEFUEL is displayed in the PRESEL window. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-30
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
Fuel Door Switch FUEL QUANTITY COMPUTER REFUEL / DEFUEL PANEL ON
L ENG
Defuel/Isolation Transfer Shutoff Valve
CENTER
RIGHT
1ØØØ
ØØØØ
2ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
Left Wing Refuel Solenoid Shutoff Valve
LEFT
LEFT
CENTER
RIGHT OPEN
DEFUEL
AUTO
DC ESS BUS 2 DECR
REFUEL
DC EMER BUS
Right Wing Refuel Solenoid Shutoff Valve
R ENG
START
CLOSE
MANUAL
DC ESS BUS 1
28 VDC
PRESEL INCR
MANUAL STOP/SOV TEST
Refuel/Defuel Adapter
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
L SURGE TANK
R MAIN TANK
COLLECTOR TANK
COLLECTOR TANK
R SURGE TANK
SOL
M
Left Wing Refuel Shutoff Valve
Manual Defuel Valve
LEGEND Center Tank Refuel Shutoff Valve Manual Defuel Center Tank Refuel Right Wing Refuel Valve Lever Solenoid Shutoff Valve Shutoff Valve
M PS SOL
Motor Shutoff Valve Pressure Switch Solenoid Shutoff Valve
CS1_CS3_2823_027
M
Figure 15: Refuel/Defuel System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-31
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUELING Auto Refueling With the AUTO/MANUAL switch in AUTO, the fuel quantity to be delivered is preselected on the RDCP using the PRESEL INCR/DECR switch setting. The REFUEL START/STOP/SOV TEST switch has a momentary STOP/TEST SOV, a START position, and a center OFF position. The START position is used for automatic refueling. When refueling is in progress in AUTO MODE, selecting STOP/SOV TEST stops refueling. The refueling operation starts on selection of the REFUEL/DEFUEL switch to REFUEL, and the START/STOP/SOV TEST switch to the START position. The refueling operation only stops when the preselected fuel quantity is delivered, or the operation is stopped by selecting the START/STOP/SOV TEST switch to the STOP/SOV TEST position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-32
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ
2ØØØ
TOTAL
PRESEL
OFF
Left Wing Refuel Solenoid Shutoff Valve
LEFT
L ENG
CENTER
RIGHT OPEN
DEFUEL
AUTO
PRESEL INCR
DECR
START
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Refuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
R SURGE TANK
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
SOL
M
Left Wing Refuel Shutoff Valve
Center Tank Refuel Shutoff Valve APU Manual Defuel Center Tank Refuel Right Wing Refuel Valve Lever Solenoid Shutoff Valve Shutoff Valve
CS1_CS3_2823_003
AC
Figure 16: Auto Refueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-33
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
Manual Refueling The manual mode is entered by selecting the MANUAL on the RDCP and the REFUEL/DEFUEL switch to the REFUEL position. MANUAL is displayed in the preselect window. The refuel SOVs are opened by setting the MANUAL REFUEL VALVE switches to OPEN. The individual tank quantities are monitored on the RDCP. When the desired tank quantity is reached, the respective refuel SOVs are closed manually using the MANUAL REFUEL VALVE switches on the RDCP. In the manual mode, refueling continues until the MANUAL REFUEL VALVE switches are closed, or the fuel high-level condition is reached. At that point, the refuel valves are closed by the FQC. These switches are hardwired to the FQC to provide a discrete to open or close their respective valve in case the RDCP fails.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-34
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ OFF
Left Wing Refuel Solenoid Shutoff Valve
MANUAL PRESEL
TOTAL LEFT
L ENG
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
DECR
START
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Refuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
R SURGE TANK
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
SOL
M
Left Wing Refuel Shutoff Valve
Center Tank Refuel Shutoff Valve APU Manual Defuel Center Tank Refuel Right Wing Refuel Valve Lever Solenoid Shutoff Valve Shutoff Valve
CS1_CS3_2823_004
AC
Figure 17: Manual Refueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-35
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
DEFUELING Suction Defueling Suction defueling is carried out using the fuel truck suction pressure. With the RDCP and refuel adapter configured for defueling, the fuel truck pumps are turned on. The maximum suction pressure allowed is -8 psig. Fuel from the left and right main tank is suctioned out through the inlet snorkel of the AC boost pumps. Fuel in the center tank is removed by opening the manual defuel valve on the rear spar of the center tank. Fuel is drawn out of the center tank through the defuel valve using fuel truck suction pressure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-36
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ OFF
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Defuel/Isolation Transfer Shutoff Valve
DEFUEL PRESEL
TOTAL LEFT
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
DECR
START
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Defuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
L SURGE TANK
COLLECTOR TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
R SURGE TANK
SOL
M
Manual Defuel Valve
APU
Manual Defuel Center Tank Refuel Valve Lever Solenoid Shutoff Valve
CS1_CS3_2823_014
AC
Figure 18: Suction Defueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-37
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
Pressure Defueling Pressure defueling is carried out using the aircraft AC boost pumps. With the RDCP and refuel adapter configured for defueling, the AC boost pumps are turned on manually. The fuel is pumped out of the main tanks through the engine feed manifold to the refuel manifold. If there is fuel in the center fuel tank, the AC boost pumps provide the motive flow to operate the transfer pumps. The transfer pumps empty the center tank fuel into the main tanks, where it is then pumped out.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-38
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
ON
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ OFF
FUEL PANEL
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Defuel/Isolation Transfer Shutoff Valve
DEFUEL PRESEL
TOTAL LEFT
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
DECR
START
Engine Feed Manifold
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Defuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
L SURGE TANK
COLLECTOR TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
R SURGE TANK
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
SOL
M
APU
Manual Defuel Center Tank Refuel Valve Lever Solenoid Shutoff Valve
Refuel Manifold
CS1_CS3_2823_005
AC
Figure 19: Pressure Defueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-39
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
COMPONENT LOCATION The following components are installed in the pressure refuel/defuel system: •
Refuel adapter
•
Refuel/defuel control panel
•
Refuel shutoff valve
•
Refuel control solenoid valve
•
Defuel/isolation transfer shutoff valve and actuator
•
Manual defuel valve and lever
REFUEL ADAPTER The adapter is located on the right wing outboard of the engine at RIB 13R. An optional refuel adapter is available and installed at RIB 13L. REFUEL/DEFUEL CONTROL PANEL The refuel/defuel control panel (RDCP) is located on the fuselage near the inboard right wing leading edge.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-40
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
Refuel/Defuel Control Panel
CS1_CS3_2823_026
Refuel Adapter
Figure 20: Refuel Adapter and Refuel/Defuel Control Panel Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-41
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL SHUTOFF VALVE The refuel shutoff valves for each main tank are mounted on the inboard side of RIB 12 inside the fuel tank. The center tank refuel shutoff valve is mounted on the inboard side of RIB 1L inside the center fuel tank. REFUEL CONTROL SOLENOID VALVE The main tank refuel control solenoid valves are mounted on a bracket at RIB 12 on the front spar. The center tank refuel control solenoid valve is mounted on the aft spar at RIB 1L.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-42
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
RIB 1L
RIB 12R
RIB 12L
Front Spar RIB 12L
Refuel Control Solenoid Valve
Refuel Control Solenoid Valve
CENTER WING BOX
LEFT WING SHOWN (RIGHT WING SIMILAR)
CS1_CS3_2823_006
Refuel Shutoff Valve
Figure 21: Refuel Control Solenoid Valve and Refuel Shutoff Valve Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-43
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
DEFUEL/ISOLATION TRANSFER SHUTOFF VALVE AND ACTUATOR The valve body is mounted inside the center tank between RIB 0 and RIB 1L on the rear spar. The 28 VDC motor actuator is mounted outside of the tank. MANUAL DEFUEL SHUTOFF VALVE AND LEVER The center tank manual defuel shutoff valve is mounted inside the center tank on the rear spar between RIB 1L and RIB 2L. A manual lever, mounted on the outside of the tank, is used to operate the valve. The valve lever is pushed in to move it open or closed. The lever assembly and valve body can be replaced independent of one another.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-44
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
Defuel/Isolation Transfer Valve
B A
RIB 0
RIB 1L
Defuel/Isolation Transfer Valve Actuator
A MANUAL DEFUEL VALVE AND LEVER
CS1_CS3_2823_025
B
Figure 22: Defuel System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-45
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
COMPONENT INFORMATION REFUEL ADAPTER The refuel adapter provides a connection point for single-point defueling/refueling of the aircraft. The refuel adapter has a poppet assembly and a fuel cap attached to the adapter by a cable assembly. During defueling, a locking mechanism is rotated to the DEFUEL position, allowing fuel to flow from the refuel manifold to the truck.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-46
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel 28-23 Refuel/Defuel System
Refuel/Defuel Selector
Refuel Adapter
CS1_CS3_2823_008
FUEL CAP REMOVED
Figure 23: Refuel Adapter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-47
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL/DEFUEL CONTROL PANEL The RDCP controls and monitors aircraft refueling and defueling. The RDCP functions are enabled when the aircraft is on the ground and the RDCP refueling door is in the open position. The refuel door switch signals the fuel quantity computer (FQC) that the refuel door is open.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-48
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel
Refuel Door Switch
Refuel Door Target
CS1_CS3_2823_011
28-23 Refuel/Defuel System
Figure 24: Refuel/Defuel Control Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-49
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel 28-23 Refuel/Defuel System
CONTROLS AND INDICATIONS REFUEL/DEFUEL CONTROL PANEL The refuel/defuel control panel (RDCP) controls and monitors the aircraft refueling and defueling operations. An indicator displays the left, center, and right fuel tank quantities, the total fuel quantity, and the load preselect quantity. The panel has switches to control power, automatic, or manual modes, and refuel and defuel modes.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-50
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ
5ØØØ
TOTAL
PRESEL
OFF
LEFT
CENTER
RIGHT OPEN
DEFUEL
AUTO
PRESEL INCR
DECR
START
CLOSE REFUEL
MANUAL STOP/SOV TEST CS1_CS3_2823_023
MANUAL
Figure 25: Refuel/Defuel Control Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-51
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
RDCP Messages
RS-422 COMMUNICATION ERROR
Messages related to refueling and defueling display on the RDCP.
A fault has occurred in the communications bus and the RDCP is inoperative. The displays remain dashed and the PRESEL portion of the window COMM ERR is displayed. The RDCP does not respond to any switch inputs.
INHIBIT When refueling the aircraft in manual mode or defueling the aircraft, an INHIBIT message is displayed whenever the RCDP is powered in manual mode or when switched to manual mode and any of the MANUAL LEFT/CENTER/RIGHT switches are in the OPEN position. A CLOSE ALL VALVES message is also displayed on the panel. The message clears when all of the MANUAL LEFT/CENTER/RIGHT switches are closed.
RCDP SWITCH FAULT The RDCP detects a switch fault (a momentary switch is held in the same position for more than 3 minutes). The message SWCH FLT will be displayed on the preselect window. The fuel tank quantities will continue to be displayed on the RDCP.
FULL
RDCP FAULT
A FULL message indicates the level in the respective tank is between 97% and 98% of total capacity. The message remains as long as the condition exists.
For other RDCP faults, the message RDCP FLT is displayed on the preselect window. If possible, the RDCP continues to display the fuel tank quantities. The quantities displayed must be verified with information from the flight deck.
HIGH LVL A HIGH LVL message indicates the level in the respective tank is greater than 98% of total capacity. The message remains as long as the condition exists.
FQC FAULT An FQC fault that affects the refueling process is detected, the message FQC FLT is displayed on the preselect window. If possible, the RDCP continues to display the quantities.
IMBAL During refueling, an IMBAL message is displayed, flashing alternately with the quantity information when there is an imbalance between the left and right tank fuel quantities. A fuel quantity difference of 364 kg (800 lb) or greater between the wing fuel tanks is considered an imbalance. An IMBAL message is displayed for the fuel tank containing the lesser of the two imbalance quantities.
AUTOMATIC REFUELING STOPPED An automatic refuel has been stopped before completion (START/STOP switch toggled to the STOP position during an automatic refuel), the STOPPED message is displayed in the preselect window for 5 seconds.
RCDP CPU FAULT A fault has occurred in the RDCP CPU. The displays remain blank and the RDCP does not respond to any switch inputs. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-52
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
CLOSE
ALL
4ØØØ
RIGHT
REFUEL / DEFUEL PANEL LEFT
CENTER
VALVES
FULL
INHIBIT
61ØØ
TOTAL
PRESEL
2000
HIGH LVL
PRESEL
CENTER
CENTER
RIGHT
1000
CENTER
RIGHT
TOTAL
PRESEL
REFUEL / DEFUEL PANEL LEFT
CENTER
----
----
4ØØØ
10000
----
----
TOTAL
PRESEL
PRESEL
CPU FAULT
FUEL TANK IMBALANCE
REFUEL / DEFUEL PANEL LEFT
LEFT
2000
2000
5ØØØ
SWC H F LT
TOTAL
PRESEL
1000
TOTAL
CENTER
---COMM ERR PRESEL
REFUEL / DEFUEL PANEL
RIGHT
LEFT
2000
2000
5ØØØ
R D C P F LT
5ØØØ
TOTAL
PRESEL
TOTAL
SWITCH FAULT MESSAGE
RIGHT
RS-422 COMMUNICATION ERROR
REFUEL / DEFUEL PANEL
RIGHT
2000
CENTER
2000 10000
2000 TOTAL
RIGHT
FUEL TANK HIGH LEVEL
REFUEL / DEFUEL PANEL LEFT
1000
61ØØ
FUEL TANK FULL
REFUEL / DEFUEL PANEL
IMBAL
LEFT
10000
TOTAL
INHIBIT
LEFT
1000
REFUEL / DEFUEL PANEL
RIGHT
1000
CENTER
1000
RIGHT
2000 F Q C F LT PRESEL
FQC FAULT
RDCP FAULT
REFUEL / DEFUEL PANEL CENTER
2000
RIGHT
1000
2000
5ØØØ
STOPPED
TOTAL
PRESEL
AUTOMATIC REFUELING STOPPED
CS1_CS3_2823_017
LEFT
Figure 26: Refuel/Defuel Control Panel Messages Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-53
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
OPTIONAL VIRTUAL REFUEL PANEL The virtual refuel panel is displayed on the FUEL synoptic page. The virtual refuel panel is displayed when the aircraft is on ground only with no engines running. To refuel the aircraft using the virtual refuel panel, the refuel door must be open, the RDCP powered, and the REFUEL switch set to AUTO. This panel allows the crew to enter the total fuel quantity in the preselect data entry field. Pressing ENTER sends the data to the FQC. If the preselect value entered exceeds the tank capacity, maximum allowable fuel quantity is assumed by the FQC. If refueling is in progress, the preselect fuel quantity can be modified. If the preselect value is less than the current fuel load, refueling stops at the current fuel quantity level. When the virtual refuel panel is in control of the preselect quantity, the RDCP ignores inputs on the RDCP PRESEL INCR/DECR switch. The refuel SOVs cannot be opened from the flight deck. The START switch on the RDCP must be selected to open the aircraft refuel SOVs. The virtual refuel panel has priority over the RDCP commands. If no value or 0 is entered in the preselect field of the virtual refuel panel, the preselect fuel quantity can be entered in the RDCP. Cycling the RDCP power for 10 seconds overrides the flight deck preselect field. The source of the preselect fuel quantity used by the FQC is displayed on the virtual refuel panel and on the RDCP. The left, center, right, and total fuel quantities are still displayed on the RDCP in their respective windows. The PRESEL window displays the message COCKPIT alternating with the preselect quantity that is entered in the flight deck refuel panel. The COCKPIT message indicates that the entry has been made in the flight deck and that refueling may be started.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-54
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
3ØØØ
43 Ø Ø Ø
3ØØØ
ON
10000 KG 0 KG
1 Ø3 Ø Ø Ø OFF
COCKPIT DECR
PRESEL
TOTAL LEFT
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
START
CLOSE
MANUAL REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
REFUEL
ACCEPTED
MANUAL STOP/SOV TEST
VALUE SET TO MAX
34000 KG IN PROGRESS
REFUEL MANAGED BY COCKPIT - COMPLETE
20 °C 3000 KG
20 °C 3000 KG
REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED 34000 KG COMPLETE
REFUEL MANAGED BY PANEL - COMPLETE
4000 KG
REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED èèèèè KG STANDBY
APU
REFUEL PRESELECT REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED 15000 KG
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED èèèèè KG COMPLETE
ACCEPTED VALUE INVALID
CS1_CS3_2823_009
REQUESTED VALUE ENTERED - SYSTEM NOT READY
Figure 27: Virtual Refuel Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-55
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
OPERATION NOTE
AUTO REFUEL OPERATION
1. If a fuel imbalance greater than 364 kg (800 lb) occurs between the wing fuel tanks, refueling must be stopped.
1. Open the REFUEL panel door. 2. Set the ON/OFF panel switch to the ON position. 3. Set the AUTO/MANUAL switch to the AUTO position. 4. Hold the PRESEL INCR/DECR switch to the INCR position to increase the desired refuel quantity displayed in the PRESEL window, or hold the switch to the DECR position to decrease the desired refuel quantity displayed in the PRESEL window.
2. When all tanks are refueled simultaneously; the time to completely refuel the center tank is approximately 30 minutes and 20 minutes to completely refuel the wing tanks. 3. If the refueling operation is stopped by selecting the START/STOP/SOV TEST to STOP/SOV TEST, a STOPPED message is displayed for 5 seconds.
5. Perform the refuel valve SOV TEST. The STOP/SOV TEST position is used to test the SOV operation at the start of refueling. The refuel valves are tested by applying fuel pressure and then selecting the SOV TEST position. An SOV TEST IN PROGRESS message appears in the display windows, alternating every 2 seconds with the fuel quantity display. The fuel valves close and an SOV TEST PASSED message is displayed on the RDCP indicating a successful test. The message is removed when any RDCP switch is selected. 6. Set the START/STOP/SOV TEST switch to the START position. The PRESEL window then displays a clock icon and the preselected fuel quantity message. 7. The refuel flow to the aircraft stops automatically when the preselected quantity is achieved. In the PRESEL window COMPLETE is displayed, alternating every 2 seconds with the preselected fuel quantity. 8. Set the ON/OFF panel switch to the OFF position and close the REFUEL panel door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-56
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL ON
LEFT
CENTER
1ØØØ
ØØØØ
REFUEL / DEFUEL PANEL
RIGHT
1ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL RESEL
OFF
LEFT
CENTER
RIGHT
DEFUEL
AUTO
PRESEL INCR
LEFT
CENTER
PASS
PASS
ON
SOV TEST
TOTAL
PRESEL
LEFT
START
PASS
2ØØØ OFF
DECR
RIGHT
CENTER
RIGHT OPEN
DEFUEL
AUTO
PRESEL INCR
DECR
START
CLOSE MANUAL STOP/SOV TO T TEST
MANUAL
REFUEL / DEFUEL PANEL LEFT
CENTER
1ØØØ
ØØØØ
REFUEL / DEFUEL PANEL
RIGHT
1ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
OPEN
DEFUEL
AUTO
PRESEL INCR
LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ
5ØØØ
TOTAL
PRESEL
OFF
DECR
LEFT
START
CENTER
SOV TEST
in
MANUAL STOP/SOV MANUA TEST
RIGHT
PROGRESS
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
LEFT
CENTER
RIGHT OPEN
DECR
CLOSE REFUEL
DEFUEL
AUTO
REFUEL
MANUAL
REFUEL / DEFUEL PANEL LEFT
ON
PRESEL INCR
START
CLOSE
MANUAL
MANUAL STOP/SOV TEST
MANUAL MANUA STOP/SOV TEST
REFUEL / DEFUEL PANEL
PRESEL INCR
LEFT
CENTER
RIGHT
25ØØ
ØØØØ
25ØØ
ON
DECR
5ØØØ
COMPLETE
TOTAL
PRESEL
OFF
LEFT
START
CENTER
RIGHT OPEN
MANUAL STOP/SOV MANUA TEST
DEFUEL
AUTO
PRESEL INCR
DECR
START
MANUAL A STOP/SOV TEST
CS1_CS3_2823_012
ON
REFUEL
MANUAL
Figure 28: Automatic Refuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-57
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
MANUAL REFUEL OPERATION 1. Open the REFUEL panel door. 2. Set the ON/OFF panel switch to the ON position. 3. Set the MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 4. Set the REFUEL/DEFUEL switch to REFUEL. 5. Set the AUTO/MANUAL switch to the MANUAL position. 6. Set the MANUAL LEFT/CENTER/RIGHT switches to the OPEN position for the tanks that are to be filled. 7. Set the MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position when the required fuel level is reached. 8. Set the ON/OFF panel switch to the OFF position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-58
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL ON
LEFT
CENTER
1ØØØ
ØØØØ
RIGHT
1ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
LEFT
CENTER
RIGHT OPEN
AUTO
DEFUEL
PRESEL INCR
DECR
START
CLOSE E
REFUEL / DEFUEL PANEL
MANUAL STOP/SOV TEST
REFUEL
MANUAL UAL
LEFT
CENTER
3ØØØ
3ØØØ
ON
9ØØØ OFF
CENTER
PRESEL RIGHT OPEN
CENTER
2ØØØ
1ØØØ
ON
5ØØØ OFF
CENTER
RIGHT
2ØØØ PRESEL
RIGHT OPEN
MANUAL
PRESEL INCR
START
REFUEL
MANUAL STOP/SOV TEST
DEFUEL
AUTO
DECR
START
REFUEL
CS1_CS3_2823_013
CLOSE
MANUAL
AUTO
MANUAL
TOTAL LEFT
DEFUEL
DECR
CLOSE
REFUEL / DEFUEL PANEL LEFT
3ØØØ
PRESEL INCR
MANUAL
TOTAL LEFT
RIGHT
MANUAL STOP/SOV TEST
Figure 29: Manual Refuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-59
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
PRESSURE DEFUEL OPERATION 1. Connect external AC power to the aircraft. 2. Open the REFUEL panel door. 3. Set the ON/OFF switch to the ON position. 4. Set all MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 5. Set the MANUAL REFUEL/DEFUEL switch to DEFUEL. Verify that the defuel message is displayed on the RDCP window. 6. On the pressure refuel adapter, set the REFUEL/DEFUEL selector to DEFUEL. 7. Connect the refuel hose nozzle to the refuel adapter. 8. Set all MANUAL LEFT/CENTER/RIGHT switches to the OPEN position. 9. Set the GRAV XFR PBA to the CLOSED position. 10. Set the BOOST PUMP switches to the ON position for the fuel tank(s) to be defueled. 11. When an AC boost pump low-pressure indication is displayed on the FUEL synoptic page, set the BOOST PUMP switch to the OFF position. 12. Set all MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 13. Set the ON/OFF switch to the OFF position. 14. Disconnect the refuel hose nozzle from the refuel adapter and turn the pressure refuel adapter REFUEL/DEFUEL selector to DEFUEL.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-60
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
FUEL MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
AUTO
OFF
ON
AUTO
OFF
R CTR
ON
GRAV XFR
REFUEL / DEFUEL PANEL
ON
LEFT
CENTER
RIGHT
2ØØØ
2ØØØ
2ØØØ
ON
FUEL PANEL
6ØØØ Refuel/Defuel Selector
OFF
DEFUEL PRESEL
TOTAL LEFT
CENTER
PRESEL INCR
RIGHT OPEN
DEFUEL
AUTO
DECR
START
CLOSE REFUEL
MANUAL STOP/SOV TEST
CS1_CS3_2823_024
MANUAL
(FUEL CAP REMOVED)
Figure 30: Pressure Defuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-61
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
SUCTION DEFUEL OPERATION 1. Open the REFUEL panel door. 2. Set the ON/OFF switch to the OFF position. 3. Set each of the MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 4. If the center fuel tank is to be defueled, the MANUAL DEFUEL valve is opened by pushing and turning the MANUAL DEFUEL LEVER, located on the rear spar of the center fuel tank. 5. On the pressure refuel adapter, set the REFUEL /DEFUEL selector to DEFUEL. 6. Set the REFUEL/DEFUEL switch to the DEFUEL position. 7. Set the POWER SWITCH to the ON position. The PRESEL display window shows DEFUEL. 8. Set the AUTO/MANUAL switch to the MANUAL position. 9. Attach the hose to the refuel adapter. Start the fuel truck pump, which must be configured in the suction mode. 10. Set all MANUAL LEFT/CENTER/RIGHT switches to the OPEN position. 11. Fuel is suctioned out of the wing tanks. 12. Monitor the LEFT/CENTER/RIGHT display window and stop the fuel truck pump when desired quantity is reached. 13. Manually close center tank MANUAL DEFUEL valve when the desired quantity in the center wing tank is reached. 14. Set the MANUAL LEFT/RIGHT switch(es) to the CLOSE position to stop defueling in all tanks. 15. Set the ON/OFF switch to the OFF position. 16. Disconnect the refuel hose nozzle from the refuel adapter and turn the pressure refuel adapter REFUEL/DEFUEL selector to DEFUEL. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-62
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
2ØØØ
2ØØØ
ON
6ØØØ OFF
DEFUEL
CENTER
DECR
PRESEL
TOTAL LEFT
PRESEL INCR
RIGHT OPEN
DEFUEL
AUTO
START
CLOSE
MANUAL
REFUEL
MANUAL STOP/SOV TEST
CS1_CS3_2823_015
MANUAL DEFUEL VALVE
Figure 31: Suction Defuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-63
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
DETAILED DESCRIPTION REFUEL POWER The fuel system has two power modes for normal and battery only refuel/defuel operation. In normal operation, the refuel panel door is closed and the system is powered by DC ESS BUS 1 and DC ESS BUS 2. When the aircraft is refueled, the refuel door is open. The refuel door switch signals the FQC that the door is opened. When the RDCP power switch is selected on, the FQC receives a 28 VDC signal indicating the refuel panel is turned on. The refuel door switch also energizes relays in CDC 1 and CDC 2 to switch the fuel power from the DC ESS BUSES to the DC EMER BUSES.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-64
CS130_28_L3_RefuelDefuelSyst_DetailedDescription.fm
28 - Fuel 28-23 Refuel/Defuel System
LEFT REFUEL SHUTOFF VALVE
CDC 1-7-21 L Fuel Qty DC ESS BUS 1
CDC 1
CENTER REFUEL SHUTOFF VALVE
RIGHT REFUEL SHUTOFF VALVE
SSPC 5A
Logic: Always On 5
Refuel Valve CTL A
Logic: Always On DC EMER BUS
5
Refuel Valve CTL B K3
REFUEL PANEL DOOR SWITCH
FUEL QUANTITY COMPUTER (FQC)
LEFT REMOTE DATA CONCENTRATOR (RDC)
CENTER REMOTE DATA CONCENTRATOR (RDC)
Lane 2
Lane 1
Lane 2
Lane 1
SSPC 5A
Lane 2
CDC 2 Door Open
R Fuel Qty DC ESS BUS 2
Channel A
CDC 2-7-20
Lane 1
K3 Channel B
DC EMER BUS
RIGHT REMOTE DATA CONCENTRATOR (RDC)
DEFUEL/ISOLATION TRANSFER SHUTOFF VALVE
REFUEL/DEFUEL CONTROL PANEL
On
Refuel Power On
CS1_CS3_2823_019
Off
Figure 32: Refuel Power (Normal Operation - Door Closed) Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-65
CS130_28_L3_RefuelDefuelSyst_DetailedDescription.fm
28 - Fuel 28-23 Refuel/Defuel System
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM The following refuel/defuel components are tested through the onboard maintenance system (OMS). Refuel Solenoid Command Test When the OMS commands a refuel solenoid command test, the selected FQC channel energizes each refuel solenoid and verifies that the solenoid is energized. The FQC channel then de-energizes each refuel solenoid and verifies that the solenoid is de-energized. This test checks the electrical command of the refuel solenoids, but does not verify fuel flow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-66
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-23 Refuel/Defuel System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-REFUEL SOLENOID COMMAND TEST
001 /004 REFUEL SOLENOID COMMAND TEST
*** CAUTION *** KEEP A/C FUEL PUMPS OFF
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2823_016
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 33: Refuel Solenoid Command Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-67
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-23 Refuel/Defuel System
Defuel Shutoff Valve Command Test When the OMS commands a defuel shutoff valve command test, the selected FQC channel commands the DEFUEL SOV open, and verifies that the SOV position feedback has reported that the valve has opened. The FQC closes the defuel SOV and verifies that the SOV position feedback reports that the valve has closed. This test checks the electrical command of the defuel shutoff valve, but does not verify fuel flow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-68
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-23 Refuel/Defuel System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-DEFUEL SHUTOFF VALVE COMMAND TEST
001 /004 DEFUEL SHUTOFF VALVE COMMAND TEST
*** CAUTION *** KEEP A/C FUEL PUMPS OFF
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2823_021
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 34: Defuel Shutoff Valve Command Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-69
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-21 Engine Feed System
28-21 ENGINE FEED SYSTEM GENERAL DESCRIPTION The feed system consists of one engine feed ejector pump and one AC boost pump in the collector tank. Each engine has its own feed system. The two engine feeds are connected to allow either AC boost pump to feed both engines. Check valves in the engine feed lines prevent a loss of fuel pressure when an ejector feed pump or boost pump fails. Normally-open engine feed shutoff valves in each engine fuel feed line stop the flow of fuel in the event of an engine fire. Each engine feed system is monitored by two pressure switches. Low fuel pressure at the engine is detected by a pressure switch in the engine feed line. Each AC boost pump has a pressure switch at the pump outlet to detect low-pressure. In the event of a total failure of the engine feed pumps, the engine can suction feed through the AC boost pump inlet up to 25,000 ft. The engine-driven pump (EDP) suctions fuel through the engine feed ejector induced port and the AC boost pump inlet snorkel. Motive flow fuel generated by the EDP powers the onside engine feed and scavenge ejector pumps. A check valve in the motive flow line isolates the engine from the fuel tank. The engine-feed ejector pump is the main source of fuel for engine operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-70
CS130_28_L2_EFS_General.fm
28 - Fuel 28-21 Engine Feed System
L ENG
R ENG
SOL
L DRY BAY
CENTER TANK
PS
SOL
R DRY BAY PS PS
AC
AC
PS
M
M M
L MAIN TANK
LEGEND
AC
COLLECTOR TANK
L SURGE TANK
Engine Feed Motive Flow AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
R SURGE TANK
M
SOL
M
One-Way Check Valve PS
Pressure Switch
Engine Feed Ejector Pump
APU
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
CS1_CS3_2821_003
Motor Shutoff Valve
M
SOL
R MAIN TANK
Figure 35: Engine Feed System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-71
CS130_28_L2_EFS_General.fm
28 - Fuel 28-21 Engine Feed System
COMPONENT LOCATION The following components are installed in the engine fuel feed system: •
Engine feed ejector pump
•
AC boost pump
•
Engine feed shutoff valve and actuator
•
Engine feed pressure switch
ENGINE FEED EJECTOR PUMP The engine feed ejector pump is located in the collector tank outboard of RIB 5. AC BOOST PUMP An AC boost pump is mounted on RIB 6 in each collector tank. ENGINE FEED SHUTOFF VALVE AND ACTUATOR The engine feed shutoff valve and actuator are installed between RIB 8 and RIB 9. ENGINE FEED PRESSURE SWITCH The engine feed pressure switch is mounted on the rear spar, near RIB 8.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-72
CS130_28_L2_EFS_CompLocation.fm
28 - Fuel 28-21 Engine Feed System
RIB 5L Engine Feed Ejector Pump Rear Spar
Engine Feed Shutoff Valve AC Boost Pump
Engine Feed Pressure Switch
Engine Feed Shutoff Valve Actuator
NOTE
CS1_CS3_2821_018
RIB 9L
AC Boost Pump Pressure Switch
Left wing shown, right wing similar.
Figure 36: Engine Feed Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-73
CS130_28_L2_EFS_CompLocation.fm
28 - Fuel 28-21 Engine Feed System
COMPONENT INFORMATION AC BOOST PUMP An AC boost pump is mounted on RIB 6 in each collector tank. The AC boost pump consists of a cannister and cartridge elements. The cannister, which contains an impeller, and suction and outlet ports, is located inside the collector tank. The suction port is connected to a pickup fitted with an inlet strainer. The pump outlet is connected to the engine feed lines. The cartridge element is mounted in the dry bay. The cartridge element has a 115 VAC, variable frequency, 3-phase motor and pressure switch. The motor drives the impeller and the pressure switch monitors the pump discharge pressure. The cannister has a built-in check valve that allows the cartridge to be removed without draining the fuel tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-74
CS130_28_L2_EFS_CompInfo.fm
28 - Fuel 28-21 Engine Feed System
RIB 5L Collector Tank RIB 5L
Pump Outlet to Engine Fuel Shutoff Valve
RIB 7L
Pump Pickup
Crossfeed Line
Dry Bay
AC Boost Pump Cannister NOTE Pressure Switch
Left wing shown, right wing similar.
CS1_CS3_2821_006
AC Boost Pump Cartridge
Figure 37: AC Boost Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-75
CS130_28_L2_EFS_CompInfo.fm
28 - Fuel 28-21 Engine Feed System
CONTROLS AND INDICATIONS ENGINE FIRE PBA The L or R ENG FIRE PBA closes the engine fuel shutoff valve. FUEL PANEL The AC L or R BOOST PUMP rotary switches are on the FUEL panel. The FUEL panel is located on the overhead panel.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-76
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
A
B
L ENG
R ENG
F IRE FIRE E
F IIRE RE E
C
BTL 1
BTL 2
BTL 1
BTL 2
AVAIL
AVAIL
AVAIL
AVAIL
OVERHEAD
A
LEFT ENGINE FIRE PBA
B
RIGHT ENGINE FIRE PBA
FUEL MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
CTR
ON
GRAV XFR
C FUEL PANEL
CS1_CS3_2821_008
ON
Figure 38: Engine Fire PBAs and Engine Fuel Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-77
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
SYNOPTIC DISPLAY The engine fuel feed can be monitored on the FUEL synoptic page. The filter shown on the display is the main engine fuel filter and is covered in ATA 73.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-78
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
VALVE POSITION NORMAL OPERATION
9000 KG 3000 KG
Symbol
Condition
VALVE POSITION FAIL OPERATION Symbol
Open
Open
Closed
Closed
In Transition
In Transition
Invalid
Invalid
PUMPS
ENGINE FILTER Symbol
Symbol
23 °C 2000 KG
23 °C 2000 KG
Condition
Condition
Condition
Active
Engine ON - Normal
Off
Engine OFF - Normal
Failed
Engine ON or OFF - Degraded
5000 KG
Invalid Invalid Engine ON - Bypass
CS1_CS3_2821_009
APU
SYNOPTIC PAGE - FUEL
Figure 39: Fuel Synoptic Page Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-79
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
DETAILED DESCRIPTION LEFT ENGINE FEED The left AC boost pump can be turned on automatically or manually using the L BOOST PUMP switch on the fuel panel. The ON or OFF positions provide manual control. When the switch is in the AUTO position, the AC boost pump is controlled by the logic processed in the control and distribution cabinet (CDC). The left AC boost pump is turned on automatically when: •
APU operation on the ground
•
Fuel imbalance requiring transfer to the right wing
•
Left engine start in flight
•
Low-fuel pressure at either engine
The pump turns off automatically once the engine is running and the ejector pump is supplying pressure. The left AC boost pump is powered through the essential bus contactor located in the electronic power center (EPC) 3. If AC electrical power is lost, the ram air turbine (RAT) powers the AC ESS BUS, allowing the AC boost pump to operate. If low-fuel pressure is detected by the engine fuel pressure switch, both AC boost pumps are automatically turned on. The dual-contact switch provides logic to the CDC. It also provides status information to the data concentrator unit module cabinet (DMC). The AC boost pump pressure switch monitors the pump output and reports to the DMC. The L ENG FIRE PBA closes the engine fuel shutoff valve (FSOV) when pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-80
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
ICCP
L ENGINE FEED SHUTOFF VALVE
L Fuel SOV DC EMER BUS
Open
3
Close Ground
L ENG
FIR RE E
Engine and APU Fire Panel
L BOOST PUMP AUTO
OFF
Position
DMC 1 AND DMC 2
ON
L ENGINE FUEL PRESSURE SWITCH
CDC 1-5-17 L Fuel Pump DC ESS SSPC 3A BUS 1
To CDC 1 and CDC 2
Logic: Advanced L AC BOOST PUMP PRESSURE SWITCH
EPC 3 L Fuel Pump Fuel Panel
AC ESS BUS
10 10 10
LEGEND
EPC
CDC 1 Pump Relay Position
ARINC 429 AFDX A664 Electrical Power Center
CS1_CS3_2821_010
L AC BOOST PUMP
Figure 40: Left Engine Feed Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-81
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
RIGHT ENGINE FEED The right AC boost pump can be turned on automatically or manually using the R BOOST PUMP switch on the fuel panel. The ON or OFF positions provide manual control. When the switch is in the AUTO position, the AC boost pump is controlled by the logic processed in the CDC. The right AC boost pump is turned on automatically when: •
Fuel imbalance requiring transfer to the left wing
•
Right engine start in flight
•
Low-fuel pressure at either engine
The pump turns off automatically once the engine is running and the ejector pump is supplying pressure. The right AC boost pump is powered by AC BUS 2. If low-fuel pressure is detected by the engine fuel pressure switch, both AC boost pumps are automatically turned on. The dual-contact switch provides logic to the CDC. It also provides status information to the DMC. The AC boost pump pressure switch monitors the pump output and reports to the DMC. The R ENG FIRE PBA closes the engine FSOV when pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-82
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
ICCP
R ENGINE FEED SHUTOFF VALVE
R Fuel SOV DC EMER BUS
Open
3
Close Ground
R ENG
Position DMC 1 AND DMC 2
FIR RE E
R ENGINE FUEL PRESSURE SWITCH
Engine and APU Fire Panel
R BOOST PUMP
To CDC 1 and CDC 2
AUTO
OFF
ON
R AC BOOST PUMP PRESSURE SWITCH
Fuel Panel
Logic: Advanced LEGEND ARINC 429 AFDX A664
R AC BOOST PUMP
CS1_CS3_2821_011
CDC 2-16-1,2,3 R Fuel Pump AC BUS 2 SSPC 10A
Figure 41: Right Engine Feed Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-83
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine feed system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-84
CS130_28_L2_EFS_MonitoringTests.fm
28 - Fuel 28-21 Engine Feed System
CAS MESSAGES
Table 3: STATUS Messages Table 1: CAUTION Messages
MESSAGE
MESSAGE
LOGIC
L ENG FUEL SOV Left fuel SOV failed to close/open properly. FAIL R ENG FUEL SOV Right fuel SOV failed to close/open properly. FAIL L ENG FUEL LO PRESS
Left engine feed pressure is low.
R ENG FUEL LO PRESS
Right engine feed pressure is low.
LOGIC
L BOOST PUMP ON
L boost pump is turned on via the control panel.
R BOOST PUMP ON
R boost pump is turned on via the control panel.
L BOOST PUMP OFF
L boost pump is turned off via the control panel.
R BOOST PUMP OFF
R boost pump is turned off via the control panel.
Table 4: INFO Messages Table 2: ADVISORY Messages
MESSAGE
MESSAGE
LOGIC
L BOOST PUMP FAIL
L boost pump failed to indicate high-pressure when turned on manually or automatically.
R BOOST PUMP FAIL
R boost pump failed to indicate high-pressure when turned on manually or automatically.
L FUEL EJECTOR FAIL
Left engine motive flow is faulty.
R FUEL EJECTOR Right engine motive flow is faulty. FAIL L ENG FUEL SOV CLSD
Left FUEL SOV is closed after the left engine fire switch is pressed.
R ENG FUEL SOV CLSD
Right FUEL SOV is closed after the right engine fire switch is pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
28 FUEL FAULT - Engine inlet pressure switch fault is detected. This may be ENG INLET triggered when the engine pressure switch is not high PRESS SW INOP when the fuel pump is on and providing high pressure. This may be triggered when there is a disagreement between the state of the fuel engine inlet pressure switch state as monitored by the CDCs and as monitored by the DMCs. 28 FUEL FAULT BOOST PUMP PRESS SW FAIL HI
TECHNICAL TRAINING MANUAL
For Training Purposes Only
Left and/or right boost pump pressure switch failed at high-pressure.
28-85
CS130_28_L2_EFS_MonitoringTests.fm
28 - Fuel 28-21 APU Feed System
28-21 APU FEED SYSTEM GENERAL DESCRIPTION The auxiliary power unit (APU) feed shutoff valve (SOV) isolates the APU from the fuel system. The APU feed shutoff valve opens when the APU switch is in the RUN or START position. The APU feed SOV closes when: •
Commanded OFF from the flight deck
•
APU protective shutdown occurs
•
APU fire pushbutton annunciator (PBA) is pressed
The APU is fed from the left hand engine feed line via an APU feed SOV mounted aft of the center fuel tank. The APU feed SOV is operated by a 28 VDC actuator. The APU feed SOV isolates the APU from the fuel system after APU shutdown or when commanded OFF from the flight deck. During APU ground operation, the left AC boost pump runs to provide APU feed. The arrangement of the isolation check valves between the main ejector pumps means that only the left ejector pump provides fuel to the APU while either AC pump can feed the APU. If the APU fuel burn causes an imbalance, the left and right AC pumps alternate automatically, depending on what side needs to have fuel consumed to correct imbalance.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-86
CS130_28_L2_AFS_General.fm
28 - Fuel 28-21 APU Feed System
L ENG
APU OFF
RUN
R ENG
START
PULL TURN
SOL
L DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
APU START SWITCH
SOL
R DRY BAY
M
M M
L MAIN TANK
AC
M
PS
SOL
Engine Feed Motive Flow AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
COLLECTOR TANK
COLLECTOR TANK
R SURGE TANK
M
M
SOL
APU
CS1_CS3_2821_012
LEGEND
L SURGE TANK
R MAIN TANK
Figure 42: APU Fuel Feed System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-87
CS130_28_L2_AFS_General.fm
28 - Fuel 28-21 APU Feed System
COMPONENT LOCATION The following components are installed in the APU fuel feed system: •
APU feed SOV and actuator
•
APU fuel line shroud (refer to figure 44)
APU FEED SHUTOFF VALVE AND ACTUATOR The actuator is mounted outside of the center fuel tank between RIB 0 and RIB 1R, and can be replaced without draining the fuel tank. The valve is located inside the fuel tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-88
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel 28-21 APU Feed System
APU Feed Shutoff Valve Body
APU Feed Shutoff Valve Actuator To APU
RIB 1R
RIB 0
CS1_CS3_2821_013
Rear Spar
Figure 43: APU Feed Shutoff Valve and Actuator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-89
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel 28-21 APU Feed System
APU FEED LINE, SHROUD, AND DRAIN MASTS The APU feed line runs through the fuselage from the center tank rear spar to the APU. The line is shrouded to collect any fuel that might leak from the line. Any fuel leak drains into the shroud and out one of two drain masts, located on the aft fuselage and the aft right side of the wingto-body fairing (WTBF). The center tank refuel shutoff solenoid valve also drains into the WTBF drain mast.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-90
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel
APU Fuel Drain Mast
APU Feed Line and Shroud
Refuel Solenoid Valve Drain
CS1_CS3_2821_014
28-21 APU Feed System
Figure 44: APU Fuel Line, Shroud and Drain Masts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-91
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel 28-21 APU Feed System
CONTROLS AND INDICATIONS APU START SWITCH The APU feed SOV is controlled from the APU start switch located on the APU panel. APU FIRE PBA The APU FIRE PBA is located on the ENGINE and APU FIRE panel.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-92
CS130_28_L2_AFS_ControlsIndications.fm
28 - Fuel 28-21 APU Feed System
A
APU
FII R F RE E
B
OVERHEAD
BTL AVAIL
APU OFF
RUN
A ENGINE AND APU FIRE PANEL
START
CS1_CS3_2821_015
PULL TURN
B APU START SWITCH
Figure 45: APU Start Switch and Fire PBA Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-93
CS130_28_L2_AFS_ControlsIndications.fm
28 - Fuel 28-21 APU Feed System
DETAILED DESCRIPTION The APU electronic control unit (ECU) controls the APU feed SOV through the relaxed contacts of the APU feed SOV relay located in the control and distribution cabinet 1 (CDC 1). When APU switch is in the OFF position, the APU feed SOV closes. In the event of an APU fire, the APU feed SOV closes. The APU feed SOV relay energizes when the fire detection and extinguishing (FIDEX) control unit detects an APU fire or when the APU FIRE PBA is pressed. Microswitches in the APU feed shutoff valve provide valve position indication on the engine indication and crew alerting system (EICAS). The APU feed SOV can be tested through the onboard maintenance system (OMS). The APU feed SOV test is listed under ATA 49 APU tests.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-94
CS130_28_L3_AFS_DetailedDescription.fm
28 - Fuel 28-21 APU Feed System
ICCP APU Fuel SOV DC EMER BUS
FIDEX CONTROL UNIT
3 Unattended Mode
APU CDC 1 FIRE
K1 APU FIRE PBA
APU FUEL SHUTOFF VALVE Close
Close Command
APU ECU
Open
Open Command
Ground
Position
EICAS SYN DMC 1 AND DMC 2
HMU
CS1_CS3_2821_016
OMS
Figure 46: APU Feed Shutoff Valve Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-95
CS130_28_L3_AFS_DetailedDescription.fm
28 - Fuel 28-21 APU Feed System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the APU feed system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-96
CS130_28_L2_AFS_MonitoringTests.fm
28 - Fuel 28-21 APU Feed System
CAS MESSAGES Table 5: CAUTION Message MESSAGE APU FUEL SOV FAIL
LOGIC APU fuel SOV failed to close/open properly.
Table 6: ADVISORY Message MESSAGE APU FUEL SOV CLSD
LOGIC APU FUEL SOV is closed after the APU fire switch is pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-97
CS130_28_L2_AFS_MonitoringTests.fm
28 - Fuel 28-22 Fuel Transfer System
28-22 FUEL TRANSFER SYSTEM GENERAL DESCRIPTION The fuel transfer system has the following subsystems: •
Scavenge system
•
Automatic transfer
•
Center tank to main tank fuel transfer
•
Manual transfer -
Main tank to main tank fuel transfer
-
Main tank to center tank fuel transfer
-
Gravity transfer
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-98
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Scavenge System
Automatic Transfer FUEL TRANSFER SYSTEM
Center to Main Tank Transfer Main Tank to Main Tank
Manual Transfer
Main Tank to Center Tank
CS1_CS3_2822_010
Gravity Transfer
Figure 47: Fuel Transfer System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-99
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
SCAVENGE SYSTEM Scavenge pumps minimize the amount of unusable fuel in the main tanks. The engine-driven pump (EDP) provides motive flow to scavenge ejector pumps. The pumps are located in the sump of each main tank and continuously feed the collector tanks.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-100
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG
SOL
L DRY BAY
CENTER TANK
PS
SOL
R DRY BAY PS PS
AC
AC
PS
M
M M
L MAIN TANK
LEGEND
AC
High-Pressure Fuel Scavenge AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
COLLECTOR TANK
R SURGE TANK
M
SOL
M
One-Way Check Valve PS
APU
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
CS1_CS3_2822_005
Motor Shutoff Valve
M
SOL
L SURGE TANK
R MAIN TANK
Figure 48: Scavenge System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-101
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
AUTOMATIC TRANSFER Automatic transfer corrects fuel imbalances when the MAN XFR switch on the FUEL panel is in the OFF position. Automatic transfer is accomplished through automatic control of the AC boost pumps. If the AC boost pumps are set to AUTO and the fuel quantity computer senses an imbalance of 182 kg (400 lb) between the left and right main tanks, the AC boost pump of the heavier tank turns on automatically. Both engines receive fuel directly from the heavier tank. When the tanks are within 45 kg (100 lb) of each other, the AC boost pump is switched off. If the fuel transfer does not correct the imbalance and the imbalance reaches 364 kg (800 lb), a FUEL IMBALANCE caution message is displayed. The FUEL synoptic page displays the operating boost pump in green, and a white crossfeed arrow indicates the direction of the crossfeed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-102
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG
STATUS FUEL TOTAL FUEL FUEL USED
SOL
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
4150 KG 6000 KG
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
2 °C 2000 KG
2 °C 2150 KG
R MAIN TANK
L MAIN TANK
0 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
M
SOL
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
APU
AC
Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
M
PS
SOL
AUTO
CTR R
ON
GRAV XFR
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
ON
FUEL PANEL
CS1_CS3_2822_007
LEGEND
ON
Figure 49: Automatic Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-103
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
CENTER TANK TO MAIN TANK FUEL TRANSFER Fuel transfer from the center tank to each main wing tank, is controlled by its corresponding transfer float valve. The transfer float valve opens when the fuel level of the main tank drops below 86% of its capacity. Fuel from the center tank is automatically transferred to the main tanks by transfer ejector pumps. The pump pickup is in the center fuel tank sump. Fuel flow only occurs if the fuel quantity in the main fuel tank is low enough that the associated transfer float valve opens to allow the fuel transfer. The motive flow for the pumps is provided by the engine feed ejector pump or the AC boost pump if it is running. When the main tank level increases above 86%, the transfer float valve closes. A check valve, in the suction port of each transfer ejector prevents fuel transfer back to the center tank. This transfer operation is fully automatic and continuously replenishes the main tanks until the center tank is empty. The automatic transfer of fuel from the center tank to the wing tanks ensures that the center tank fuel is always emptied first. The center tank fuel transfer is indicated on the synoptic page by green arrows pointing outward from the center tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-104
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG
STATUS S FUEL FUE EL TOTAL FUEL FUEL USED
SOL
AIR
ELEC
FLT CTR CTRL TRLL
HYD YD
INFO FO
CB B
9000 KG 6000 KG
SOL
L DRY BAY
R DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
M
M M
L MAIN TANK
23 °C 2000 KG
R MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 2000 KG 1000 KG
R SURGE TANK
APU
SYNOPTIC PAGE - FUEL M
SOL
M
AC
Center-to-Main Tank Transfer High-Pressure Fuel Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve
Transfer Ejector
APU
Transfer Ejector
Fuel Filter Motor Shutoff Valve
M
CS1_CS3_2822_002
LEGEND
One-Way Check Valve PS
Pressure Switch
Refuel Adapter Shutoff Valve SOL
Solenoid Shutoff Valve
Figure 50: Center Tank to Main Tank Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-105
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
MANUAL FUEL TRANSFER Fuel can also be manually transferred from wing tank to opposite wing tank or to center tank through appropriate selection of fuel switches on the integrated cockpit control panel (ICCP): •
Main tank to main tank, using FUEL MAN XFER switch
•
Main tank to the center tank, using FUEL MAN XFER switch and L BOOST PUMP or R BOOST PUMP switch
•
Main to main tank by gravity, using GRAV XFR PBA
Manual fuel transfer is accomplished by connecting the engine feed line to the refuel manifold using the defuel/isolation transfer valve and the refuel shutoff valves. The defuel/isolation transfer valve and refuel shutoff valves are controlled by the fuel quantity computer (FQC). Main Tank to Main Tank Fuel Transfer Selection of the FUEL MAN XFER switch to L or R opens the defuel/isolation transfer SOV and left or right main tank refuel SOV. When the L or R BOOST PUMP switches are in AUTO, the AC boost pump in the tank the fuel is transferred from turns on. When 182 kg (400 lb) has been transferred, an EICAS FUEL MAN XFR COMPLETE advisory message is displayed. The defuel/isolation transfer SOV and the refuel SOV close and the left or right boost pump turns off.
CAUTION Do not operate the L(R) AC boost pump if the fuel quantity in the associated tank is less than 363 kg (800 lb). The boost pump can be damaged.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-106
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Right Wing Refuel Solenoid Shutoff Valve
R ENG
Defuel/Isolation Transfer Shutoff Valve
STATUS FUEL TOTAL FUEL FUEL USED
SOL
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
4100 KG 3000 KG
SOL
L DRY BAY
R DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
M
M M 23 °C 2100 KG
R MAIN TANK
L MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 2000 KG 0 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
M
SOL
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
CTR
ON
GRAV XFR
APU
LEGEND
ON
M
PS
SOL
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve
Manual Defuel Valve Lever
Center Tank Refuel Solenoid Shutoff Valve
Right Wing Refuel Shutoff Valve
Solenoid Shutoff Valve
FUEL PANEL
CS1_CS3_2822_006
AC
Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
Figure 51: Main Tank to Main Tank Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-107
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Main Tank to Center Tank Fuel Transfer When the MAN XFR switch is set to CTR, the defuel/isolation transfer SOV and the center tank refuel SOV open. The left or right boost pump operation is not automatic and must be selected ON to enable fuel transfer. A FUEL XFR CTR READY status message is displayed if the pumps are not turned on. When 364 kg (800 lb) has been transferred, an EICAS FUEL MAN XFR COMPLETE advisory message is displayed. The defuel/isolation transfer SOV and the center tank refuel SOV close. The AC boost pump switch must be returned to the AUTO or OFF position, otherwise the pump continues to run.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-108
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Right Wing Refuel Solenoid Shutoff Valve
R ENG
STATUS FUEL TOTAL FUEL FUEL USED
SOL
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
2000 KG 0 KG
SOL
L DRY BAY
R DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
M
M M 23 °C 500 KG
R MAIN TANK
L MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 1000 KG 500 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
M
APU
SOL
L
Center Tank Refuel Shutoff Valve
LEGEND
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
CTR
ON
GRAV XFR ON
M
PS
SOL
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
Manual Defuel Center Tank Refuel Solenoid Shutoff Valve Valve Lever
FUEL PANEL
NOTE Aircraft shown on ground.
CS1_CS3_2822_009
AC
Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
R
L BOOST PUMP
Figure 52: Main Tank to Center Tank Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-109
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Gravity Transfer The GRAV XFR PBA transfers fuel between the main tanks to correct the imbalances between the wing tanks. When the GRAV XFR PBA is pressed, it opens the gravity crossflow SOV. The gravity crossflow SOV is normally closed to provide isolation between both wing tanks. The GRAV XFR PBA is hardwired to the gravity crossflow shutoff valve. When the SOV is open, the fuel level in the main tanks is allowed to equalize through gravity. The gravity crossflow is displayed on the FUEL synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-110
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG TOTAL FUEL FUEL USED
SOL
AIR
ELEC
FLT CTRL
HYD
INFO
CB
4000 KG 0 KG
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
23 °C 2000 KG
R MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 2000 KG 0 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
L
SOL
APU
AC
Gravity Transfer AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
M
PS
SOL
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
R BOOST PUMP
AUTO
OFF
LEGEND
R
L BOOST PUMP ON
AUTO
CTR R
OFF
ON
GRAV XFR
Gravity Crossflow Shutoff Valve
Gravity Feed Manifold
ON
FUEL PANEL
CS1_CS3_2822_004
M
Figure 53: Gravity Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-111
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
COMPONENT LOCATION The scavenge system consists of the following component: •
Scavenge ejector pump
The center tank to main tank transfer system consists of the following components: •
Transfer float valve
•
Transfer ejector pump
The gravity fuel transfer system consists of the following component: •
Gravity crossflow shutoff valve
SCAVENGE EJECTOR PUMP The scavenge ejector pumps are mounted on RIB 7 of each wing tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-112
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
RIB 5L Collector Tank Motive Flow Line
NOTE SCAVENGE EJECTOR PUMP
Left wing shown, right wing similar.
CS1_CS3_2821_004
RIB 7L
Figure 54: Scavenge Ejector Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-113
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
TRANSFER FLOAT VALVE This valve is attached to the center tank transfer discharge line at RIB 16. TRANSFER EJECTOR PUMP Each of the two transfer systems contains an ejector pump installed at RIB 4 of the center fuel tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-114
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
RIB 16L
Transfer Float Valve
RIB 4L
Transfer Ejector Pump
TRANSFER FLOAT VALVE
NOTE Left wing shown. Right wing similar.
TRANSFER EJECTOR PUMP
CS1_CS3_2822_015
Transfer Ejector Pump Pickup
Figure 55: Center Tank to Main Tank Fuel Transfer Ejector Pump and Float Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-115
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
GRAVITY CROSSFLOW SHUTOFF VALVE The gravity crossflow shutoff valve is located inboard of RIB 2R on the rear spar.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-116
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
Gravity Crossflow Shutoff Valve
RIB 2
Gravity Crossflow Shutoff Valve Actuator
Gravity Valve Pipe RIB 1R
CS1_CS3_2822_003
Rear Spar
Figure 56: Gravity Crossflow Shutoff Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-117
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
CONTROLS AND INDICATIONS Fuel transfer is controlled from the FUEL panel on the right inboard integrated cockpit control panel (ICCP). The MAN XFR switch is used to manually transfer fuel between the tanks. The MAN XFR switch is a four-position rotary switch located on the FUEL panel. The normal position of the switch is OFF. In the OFF position, automatic imbalance correction is enabled. The GRAV XFR PBA opens the gravity crossflow SOV. The gravity crossflow SOV is hardwired to the valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-118
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
FUEL MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
R CTR
ON
GRAV XFR ON
CS1_CS3_2823_010
FUEL PANEL
Figure 57: Fuel Transfer Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-119
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
SYNOPTIC PAGE Fuel transfer operations are shown on the FUEL synoptic page and the EICAS page. The arrows show the direction of transfer during an automatic transfer. The refuel valves are only shown in amber if they fail during a manual transfer operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-120
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
9500 KG 3000 KG
VALVE POSITION FAIL OPERATION
VALVE POSITION NORMAL OPERATION Symbol
Symbol
Condition
Condition
Symbol
Condition
Symbol
Condition
Open
Open
Active
Active
Closed
Closed
Off
Failed/Low
In Transition
In Transition
Failed
No Transfer
Invalid
Invalid
Invalid
Invalid
FLOW LINES Symbol
23 °C 2500 KG
CENTER TO WING XFERS
PUMPS
TRANSFER FUNCTION ACTIVE
Condition
Symbol
Fuel Flow
23 °C 2000 KG
Condition To R Wing
No Fuel Flow
To L Wing
Invalid Fuel Flow
5000 KG
TOTAL FUEL (KG)
APU
2500
5000 EICAS PAGE
Refuel Shutoff Valves SYNOPTIC PAGE - FUEL
9500 2000 CS1_CS3_2822_008
TOTAL FUEL FUEL USED
Figure 58: Fuel Synoptic Page Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-121
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the fuel transfer system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-122
CS130_28_L2_FTS_MonitoringTests.fm
28 - Fuel 28-22 Fuel Transfer System
CAS MESSAGES Table 9: STATUS Messages
Table 7: CAUTION Messages MESSAGE
LOGIC
MESSAGE
LOGIC
FUEL IMBALANCE
Fuel imbalance between left and right tank greater than 364 kg (800 lb).
FUEL MAN XFR TO L
Manual XFR TO L WING is selected on the fuel control panel.
FUEL CTR XFR FAIL
Fuel transfer out of center tank has failed, center tank fuel may become unusable.
FUEL MAN XFR TO R
Manual XFER TO R WING is selected on the fuel control panel.
FUEL MAN XFR FAIL
Manual transfer failed to perform the commanded operation or to stop the transfer.
FUEL MAN XFR TO CTR
MAN XFR to the center tank is selected and either L or R boost pump has been turned on.
FUEL GRAV XFR GRAVITY SOV has been commanded open via the ON overhead panel.
Table 8: ADVISORY Messages MESSAGE FUEL MAN XFR COMPLETE
LOGIC Wing-to-wing target fuel transfer of 182 kg (400 lb) or wing-to-center target fuel transfer of 364 kg (800 lb) has completed.
FUEL XFR CTR READY
MAN XFR to the center tank is selected but no pump has been selected.
FUEL GRAV XFR FUEL GRAV SOV failed to open or close. FAIL FUEL CTR XFR FAULT
Fuel transfer from center to the left or right tank has failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28-123
CS130_28_L2_FTS_MonitoringTests.fm
28 - Fuel 28-40 Fuel Indicating System
28-40 FUEL INDICATING SYSTEM GENERAL DESCRIPTION The fuel indicating system includes a fuel quantity computer (FQC), remote data concentrators (RDC), fuel probes, and temperature sensors. The purpose of the fuel probes is to measure the quantity of fuel in the tank. The probes are located strategically throughout the fuel tank to provide the total quantity of fuel in the tank. Compensator probes provide fuel dielectric value used to compensate the fuel quantity for changes in fuel temperature. Low-level fuel probes provide a warning when approximately 30 minutes fuel remain. The high-level fuel probes prevent an overfill of the fuel tanks. Temperature sensors in each main tank provide the fuel temperature input to the fuel indicating system. The main fuel tank temperatures are displayed on the FUEL synoptic page. The fuel probes and compensators in each tank are organized in two arrays called lane 1 and 2. Each array is connected to a processor in the RDC. Each RDC has two independent and identical processors that convert the fuel probe data into digital format. The RDCs send the tank data digitally to the FQC. The FQC has dual independent and identical processors for processing the data received from the RDC units. The FQC performs all of the computations associated with the fuel system. The FQC provides fuel quantity and temperature information to the engine indication and crew alerting system (EICAS). The FQC and RDC software is field-loadable.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-124
CS130_28_L2_FIS_General.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL FUEL
MAN MANXFR XFR OFF OFF
LL
REFUEL / DEFUEL PANEL
BOOSTPUMP PUMP LL BOOST OFF
OFF
AUTO AUTO
LEFT
ON
RR
CENTER
RIGHT
PRESEL INCR
BOOSTPUMP PUMP RR BOOST
ON
R CTR CTR
ON
GRAV XFR GRAV XFR
OFF
AUTO AUTO
OFF
OFF
ON
TOTAL
PRESEL
DECR
ON LEFT
CENTER
RIGHT OPEN
ON ON
DEFUEL
AUTO
START
CLOSE
MANUAL
FUEL PANEL
REFUEL
MANUAL STOP/SOV TEST
REFUEL/DEFUEL CONTROL PANEL Fuel Door Switch
FQC Primary
REFUEL SHUTOFF VALVES
Backup
DC EMER BUS DC ESS BUS 1
Fuel Quantity Computer Channel 1
EICAS
28 VDC
Fuel Quantity Computer Channel 2
DC ESS BUS 2
SYN OMS HMU
CENTER TANK
Fuel Quantity Probes Array 2
Fuel Quantity Probes Array 1
Right RDC
RIGHT MAIN TANK
RIGHT MAIN TANK TEMPERATURE SENSOR
CS1_CS3_2841_009
LEFT MAIN TANK
Fuel Quantity Probes Array 2
Remote Data Concentrator ARINC 429 CAN BUS CAN BUS RS-422
Center RDC Fuel Quantity Probes Array 1
RDC
Fuel Quantity Probes Array 2
LEGEND
Left RDC Fuel Quantity Probes Array 1
LEFT MAIN TANK TEMPERATURE SENSOR
Figure 59: Fuel Quantity Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-125
CS130_28_L2_FIS_General.fm
28 - Fuel 28-40 Fuel Indicating System
COMPONENT LOCATION The fuel quantity indication system components consist of: •
Fuel probes
•
Temperature sensor
•
Remote data concentrator (refer to figure 61)
•
Fuel quantity computer (refer to figure 62)
FUEL PROBES There are 58 fuel probes installed in the aircraft fuel tanks. TEMPERATURE SENSORS There is one temperature sensor installed in each of the aircraft main fuel tanks. The temperature sensor is mounted in a housing on RIB 7.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-126
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
LEGEND
CS1_CS3_2822_011
Temperature Sensor Lane 1 Fuel Probe Lane 2 Fuel Probe Lane 2 Fuel Probe/Compensator Lane 1 Fuel Probe/Compensator Lane 1 Fuel Probe (Derived Low Level) Lane 2 Fuel Probe (Derived Low Level) Lane 1 Fuel Probe (Derived High Level) Lane 2 Fuel Probe (Derived High Level)
Figure 60: Fuel Quantity Probes and Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-127
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
REMOTE DATA CONCENTRATOR The wing tank remote data concentrators (RDCs) are located on the front spar outboard of the engines. The center RDC is mounted in the wing-to-body fairing (WTBF) on FRAME 49.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-128
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
Center RDC
Frame 49
REMOTE DATA CONCENTRATOR (LEFT WING SHOWN, RIGHT WING SIMILAR)
CENTER REMOTE DATA CONCENTRATOR (LEFT WHEEL BIN REMOVED)
CS1_CS3_2841_003
Left Wing RDC
Figure 61: Remote Data Concentrator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-129
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL QUANTITY COMPUTER The fuel quantity computer (FQC) is located in the mid equipment bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28-130
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
CS1_CS3_2841_004
MID EQUIPMENT BAY AFT RACK
FUEL QUANTITY COMPUTER
Figure 62: Fuel Quantity Computer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-131
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL PROBES The tank units are two element concentric metal tube assemblies open at each end to allow fuel in. The probes are mounted on brackets attached to ribs and located throughout the tank. The length of each tank unit is determined by the height of the fuel tank at its particular location. In addition to providing quantity information, some fuel quantity probes perform other functions, as follows: •
High-level probes - There are two capacitance type fuel high-level probes installed in each main fuel tank and four in the center tank. The high-level probes are mounted outboard of RIB 20 in the main tanks, and at RIB 6 on the left and right sides of the center tank
•
Low-level probes - There are four capacitance type low-level fuel probes in each wing tank. Two fuel probes are located in each main tank outboard of RIB 7, and two are in each collector tank on the inboard side of RIB 7
•
Compensator probes - The compensator is a multiple concentric metal tube assembly mounted near the bottom of each main tank. In this position, the compensator is always covered with fuel until each tank almost empty. Any change in capacitance is the result of change of dielectric constant of the fuel. The compensator probes are located at RIB 5 of the main tanks and RIB 0 of the center tank
TEMPERATURE SENSOR The temperature sensor attaches to the housing with a bayonet fitting. The sensor is accessed from the dry bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-132
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL PROBE COMPENSATOR LOW-LEVEL FUEL PROBE
FUEL PROBE
TEMPERATURE SENSOR
CS1_CS3_2822_013
HIGH-LEVEL FUEL PROBE
Figure 63: Fuel Probes and Temperature Sensor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-133
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
DETAILED COMPONENT INFORMATION FUEL QUANTITY COMPUTER INTERFACE The fuel quantity computer interfaces with other aircraft systems through the data concentrator unit module cabinet (DMC). The FQC provides fuel quantity and fuel temperature information to the DMC via ARINC 429 BUSES. The FQC provides: •
Defuel and refuel valve control
•
Fuel quantity and temperature related CAS messages
•
High and low fuel level warning
•
Fuel imbalance warning
•
Automatic wing fuel transfer
•
Communication with the DMC
•
Onboard maintenance system communication
•
Communication with the RDCs
Most fuel system electrical components interface directly with the electrical system. The control signals are derived from the direct discrete signals, such as the fire handle, or indirect signals via an ARINC 429 BUS. The operational commands from these systems control the operation of the fuel system pumps and valves. The operating status of the pumps and valves are fed to the DMC and the electrical system. The fuel system components interface with the DMC to provide information to the engine indication and crew alerting system (EICAS), FUEL synoptic page display, and the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-134
CS130_28_L3_FIS_DetailedCompInfo.fm
28 - Fuel 28-40 Fuel Indicating System
EPC
FUEL
Gravity Crossflow Valve
MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
AUTO
ON
R CTR
L ENG Fuel Press
L ENG LO Pressure
R ENG Fuel Press
L ENG LO Pressure
Position
ON
Pressure
GRAV XFR ON
FUEL PANEL
APU Fire PBA
L AC Boost Pump
R AC Boost Pump Pressure
FIDEX Control Unit APU Switch On APU Emergency Stop
APU Electronic Control Unit
Relay K1 CDC 1
APU Feed Shutoff Valve DMC 1 and DMC 2
Position
L ENG Run Switch
L EEC
R ENG Run Switch
R EEC
SYN
L ENG Fire PBA
L ENG FEED SOV
OMS
EICAS
Position HMU
R ENG FEED SOV
LEGEND
REFUEL / DEFUEL PANEL LEFT
ON
Position
OFF
TOTAL LEFT
CDC EPC RDC SOV
Control and Distribution Cabinet Electrical Power Center Remote Data Concentrator Shutoff Valve A664 ARINC 429 CAN A CAN B RS-422
L Refuel SOV
RIGHT
DEFUEL
AUTO
Defuel/ Isolation Transfer SOV
PRESEL INCR
DECR
START
CLOSE
R RDC
C Refuel SOV
CENTER
RIGHT
PRESEL
OPEN
C RDC L RDC
CENTER
MANUAL
R Refuel SOV
REFUEL
STOP/SOV TEST
FUEL QUANTITY COMPUTER
Position CS1_CS3_2841_006
R ENG Fire PBA
Figure 64: Fuel Quantity Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-135
CS130_28_L3_FIS_DetailedCompInfo.fm
28 - Fuel 28-40 Fuel Indicating System
CONTROLS AND INDICATIONS The fuel quantity indication is displayed on the FUEL synoptic page. The fuel quantity, fuel used, and temperature information are also displayed on the EICAS page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-136
CS130_28_L2_FIS_ControlsIndications.fm
28 - Fuel 28-40 Fuel Indicating System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
9000 KG 0 KG
QUANTITY Symbol
23 °C 2000 KG
Condition
Symbol
3000 KG
Normal
1600 KG
Low/Overfilled/ Imbalance/Trapped
è½ KG
23 °C 2000 KG
TEMPERATURE
22 °C è44 °C è½ °C
Invalid
TOTAL FUEL (KG) 2000
5000 KG
5000
Condition Normal Low or High Invalid
9000 2000
EICAS PAGE
CS1_CS3_2841_002
APU
SYNOPTIC PAGE - FUEL
Figure 65: Fuel Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-137
CS130_28_L2_FIS_ControlsIndications.fm
28 - Fuel 28-40 Fuel Indicating System
DETAILED DESCRIPTION The fuel quantity system consists of arrays of AC capacitance probe units, located in the left wing, right wing and center fuel tanks. Each fuel tank uses an RDC to isolate the tank sensors from the dual-channel FQC. The RDCs provide information digitally to the FQC via two CAN BUSES. The FQC interfaces with a refuel/defuel control panel using an RS-422 BUS. Each RDC has two lanes for redundancy. The RDC lanes 1 and 2 have independent and identical processors that receive analog data from the fuel probes and sensors. An interprocessor connects the lanes within the RDC. This ensures that all the data from all the tank probes is available in either lane. Each lane is broken down into groups of fuel probes. An excitation generator in the RDC provides a separate excitation signal to each group. By grouping the fuel probes, the total energy in each circuit is very low. This reduces ignition source hazards. One fuel probe compensator is located in each tank. Temperature sensors provide the fuel temperature of the left and right wing tank fuel. To ensure availability of data to FQC from either wing compensator probe or temperature sensor in case of a CAN BUS failure, the temperature sensor in left tank is connected to RDC lane 2 while the temperature sensor in the right tank is connected to RDC lane 1. The compensator in the left tank is connected to RDC lane 2 and the compensator in the right tank is connected to RDC lane 1. The FQC has two identical and independent channels for processing the fuel quantity data. Each channel receives data digitally from the RDCs. The FQC uses the fuel quantity information to control the refuel/defuel operations and automatic fuel balancing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The FQC channels communicate with each other over an internal bus. The internal bus allows sharing of any input data from the other channel in case of input failures. Each FQC channel sends data to the L and R DMCs over ARINC 429 BUSES. In normal operation, the DMCs use the data from FQC channel 1. Channel 2 data is used when channel 1 fails. The FQC continuously monitors the fuel quantities in each tank and provides annunciation of potentially hazardous conditions. The low-level fuel probes in each wing tank are located at a level to provide an output at 30 minutes fuel remaining, which is approximately 443.2 kg (975 lb). The FQC provides two independent low-level indications to the DMC over ARINC 429 BUSES and discrete lines. In the event of an ARINC 429 BUS failure, the discrete signals are used for display on the EICAS for L or R FUEL LO QTY messages. The high-level fuel probes provide a signal when the fuel level exceeds 98% of the tank volume. During refueling, a high fuel level condition shuts down the refuel system and annunciates a TANK HIGH LEVEL on the refuel/defuel control panel (RDCP). The flight crew has no indication of a leak other than information derived from fuel quantity data. A FUEL LEAK SUSPECT caution message provides indication of a possible fuel leak if any of the following conditions are present: •
The fuel system suspects a wing tank overfill condition is present
•
The electronic engine control detects excessive fuel consumption by comparing both engines
•
The FMS calculated fuel used and indicated fuel onboard is different
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-138
CS130_28_L3_FIS_DetailedDescription.fm
28 - Fuel 28-40 Fuel Indicating System
FMS
DMC 1 AND DMC 2
EICAS
REFUEL/DEFUEL CONTROL PANEL
28 VDC
DC EMER BUS Fuel Quantity Computer Channel 1
FUEL QUANTITY COMPUTER
SYN
Fuel R Low Level
Fuel L Low Level
LEFT EEC AND RIGHT EEC
Fuel Quantity Computer Channel 2
DC ESS BUS 1
OMS HMU
DC ESS BUS 2
L REFUEL VALVE
CENTER TANK RDC
LEFT WING RDC
Lane 2
Lane 1
Lane 2
Lane 1
LEFT WING LANE 1 CENTER TANK LANE 1 AND COMPENSATOR
RIGHT WING LANE 1 AND COMPENSATOR
CENTER TANK LANE 2
LEGEND Remote Data Concentrator ARINC 429 CAN BUS CAN BUS Discrete Excitation RS-422
R REFUEL VALVE
RIGHT WING LANE 2
LEFT WING LANE 2 AND COMPENSATOR
RDC
Lane 2
C REFUEL VALVE
LEFT MAIN TANK TEMPERATURE SENSOR
RIGHT MAIN TANK TEMPERATURE SENSOR
CS1_CS3_2841_005
Lane 1
RIGHT WING RDC
DEFUEL/ ISOLATION TRANSFER VALVE
Figure 66: Fuel Quantity System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-139
CS130_28_L3_FIS_DetailedDescription.fm
28 - Fuel 28-40 Fuel Indicating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the fuel quantity indication system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-140
CS130_28_L2_FIS_MonitoringTests.fm
28 - Fuel 28-40 Fuel Indicating System
CAS MESSAGES
Table 12: INFO Messages Table 10: CAUTION Messages
MESSAGE
MESSAGE
28 FUEL FAULT - Fuel computer channel fault is detected. COMPUTER REDUND LOSS
LOGIC
L FUEL LO QTY
Low-level detected in left wing tank.
R FUEL LO QTY
Low-level detected in right wing tank.
FUEL TANK LO TEMP
Fuel temp is near Jet A fuel freezing point (-37 ºC).
FUEL TANK HI TEMP
Fuel temp above the operating limit.
FUEL COLLECTOR LO LVL
Fuel in the collector is low - suspect the transfer ejector is inoperative for the left or right tank.
FUEL LEAK SUSPECT
This may be caused by fuel leaking overboard due to: • A wing tank fuel overfill condition from the center transfer float valve stuck open • A manual transfer refuel solenoid valve failed to close Or this message may be displayed when: • The EEC suspects a leak at the engine based upon comparison between the L and R engine fuel consumptions • The FMS detects a suspected fuel leak triggered by a difference in fuel onboard compared to the FMS calculated fuel consumption while the engines are running
Table 11: ADVISORY Messages MESSAGE
LOGIC
FUEL Fuel quantity computer has failed or loss of both COMPUTER FAIL ARINC 429 channels from FQC to DMC. FUEL FAULT
Loss of redundant or non-critical function for the fuel system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
28 FUEL FAULT - Loss of fuel quantity in the left tank. L WING RDC INOP 28 FUEL FAULT L WING RDC REDUND LOSS
Fuel left wing remote data concentrator fault is detected.
28 FUEL FAULT R WING RDC INOP
Loss of fuel quantity in the right tank.
28 FUEL FAULT R WING RDC REDUND LOSS
Fuel right wing remote data concentrator fault is detected.
28 FUEL FAULT CTR WING RDC INOP
Loss of fuel quantity in the center tank.
28 FUEL FAULT CTR WING RDC REDUND LOSS
Fuel center tank remote data concentrator fault is detected.
28 FUEL FAULT FUEL GAUGING SNSR INOP
Fuel gauging sensor (probe/compensator) fault is detected.
28 FUEL FAULT FUELING DOOR OPEN
Fuel door is not CLSD prior to flight.
28 FUEL FAULT FUEL KG-LB MISCOMPARE
Fuel system refuel panel and cockpit primary page use different units (kg/lb).
TECHNICAL TRAINING MANUAL
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28 - Fuel 28-40 Fuel Indicating System
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-40 Fuel Indicating System
Table 12: INFO Messages MESSAGE
LOGIC
28 FUEL FAULT DEFUEL/XFR SOV INOP
DEFUEL/XFER SOV has failed when the actual reported SOV position is different from the commanded position, including conflicting reported position.
28 FUEL FAULT CONFIG STRAPPING INOP
Fuel gauging computer or fuel remote data concentrator has strapping fault(s).
28 FUEL FAULT SOFTWARE LOAD INOP
Fuel gauging computer software load fault is detected.
28 FUEL FAULT FUEL TEMP SNSR INOP
One fuel temperature sensor has failed.
28 FUEL FAULT GAUGING SNSR SHORT CIRCUIT
Gauging sensor short circuit of fuel probe or wiring.
28 FUEL FAULTBOOST PUMP PRESS SW FAIL HI
Left and/or right boost pump pressure switch is failed to high-pressure.
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TECHNICAL TRAINING MANUAL
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CS130_28_L2_FIS_MonitoringTests.fm
28 - Fuel 28-40 Fuel Indicating System
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM Each channel of the FQC is tested using the OMS. The FQC enters the OMS test mode when the aircraft is weight-on-wheels (WOW) and the parking brake is set. The OMS provides a system parameter page that displays the real time reported capacitance and temperature sensor data as well as status and other fuel system related faults from the ARINC 429 data. Low-Level Discrete Output Test When the OMS commands a low-level discrete output test, the FQC channel sets the FUEL L LOW LEVEL and then the FUEL R LOW LEVEL discrete outputs true for 3 seconds. The FQC also compares the DMC reported feedback of the FUEL L LOW LEVEL and FUEL R LOW LEVEL discrete outputs to the FQC commanded state. This test checks that the low-level discrete outputs operate properly, and verifies proper wiring and communication with the DMCs. The FQC reports the pass/fail status of this test to the OMS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28-144
CS130_28_L2_FIS_PracticalAspects.fm
28 - Fuel 28-40 Fuel Indicating System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-LOW LEVEL DISCRETE OUTPUT TEST
001 /004 FUEL LEVEL OUTPUT TEST
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2841_001
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 67: Low-Level Discrete Output Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-145
CS130_28_L2_FIS_PracticalAspects.fm
28 - Fuel 28-40 Fuel Indicating System
Additional Functions The fuel quantity computer provides the ability to clear non-volatile memory (NVM) faults and report the configuration from the OMS test screen selection. These initiated functions include: •
The CLEAR NVM function clears faults stored in the FQC NVM
•
The REPORT CONFIGURATION function reports the configuration data of the fuel quantity system. The FQC, RDCs, and RDCP report their configuration data to the OMS
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-40 Fuel Indicating System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-CLEAR NVM
001 /004 CLEAR NVN TEST
* NOTE * KEEP A/C FUEL PUMPS OFF
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2823_020
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 68: Clear NVM Function Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28 - Fuel 28-40 Fuel Indicating System
Page Intentionally Left Blank
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ATA 30 - Ice and Rain Protection
BD-500-1A10 BD-500-1A11
30 - Ice and Rain Protection
Table of Contents 30-81 Ice Detection System....................................................30-2 General Description .........................................................30-2 Component Location ........................................................30-4 Ice Detector..................................................................30-4 Detailed Description ........................................................30-6 Monitoring and Tests ........................................................30-8 CAS Messages ............................................................30-9 Initiated Built-In Test........................................................30-10 30-12 Wing Anti-Ice System .................................................30-12 General Description ........................................................30-12 Wing Anti-Ice Ducting ................................................30-14 Component Location .....................................................30-16 Wing Anti-Ice Valves..................................................30-16 Venturis......................................................................30-16 Telescopic Ducts........................................................30-16 Piccolo Tubes ............................................................30-16 Wing Anti-Ice Temperature Sensors..........................30-16 Pressure Sensors ......................................................30-16 Integrated Air System Controllers ..............................30-18 Controls and Indications .................................................30-20 Anti-Ice Control Panel ................................................30-20 Wing Anti-Ice indications............................................30-22 Detailed Description .......................................................30-24 Integrated Air System Controllers ..............................30-24 Wing Anti-Ice Operation.............................................30-26 Wing Anti-Ice Activation Control ................................30-28 Monitoring and Tests .....................................................30-30 System Test ...............................................................30-30 CAS Messages ..........................................................30-32 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Practical Aspects ........................................................... 30-36 Wing Anti-Ice Valve................................................... 30-36 30-22 Cowl Anti-Ice System................................................. 30-38 General Description ...................................................... 30-38 Component Location ..................................................... 30-40 Cowl Anti-Ice Valves ................................................. 30-40 Pressure Sensor ....................................................... 30-40 Fan Cowl Temperature Sensor ................................. 30-40 Swirl Nozzle .............................................................. 30-40 Detailed Component Information ................................... 30-42 Swirl Nozzle .............................................................. 30-42 Controls and Indications ................................................ 30-44 Indications ................................................................. 30-46 Detailed Description ...................................................... 30-48 Temperature and Pressure Sensing ......................... 30-48 Monitoring and Tests ..................................................... 30-50 CAS Messages ......................................................... 30-51 Practical Aspects ............................................................ 30-52 Cowl Anti-Ice Valves ................................................. 30-52 34-10 Probe Heat System.................................................... 30-54 General Description ...................................................... 30-54 Component Location ..................................................... 30-56 Angle-of-Attack Vane ................................................ 30-56 Air Data System Probe.............................................. 30-56 Total Air Temperature Probe..................................... 30-56 Static Inverter ............................................................ 30-56 Controls and Indications ................................................ 30-58 Detailed Description ..................................................... 30-60 Air Data System Probe.............................................. 30-60
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-i
CS130_M5_30_IceRainTOC.fm
30 - Ice and Rain Protection
Monitoring and Tests .....................................................30-62 CAS Messages ..........................................................30-63 30-41 Windshield and Side Window Heating System...........30-68 General Description .......................................................30-68 Component Location .....................................................30-70 Windshield .................................................................30-70 Side Window ..............................................................30-70 Windshield Ice Protection Controllers ........................30-70 Controls and Indications ................................................30-72 Detailed Description .......................................................30-74 Monitoring and Tests .....................................................30-76 CAS Messages ..........................................................30-77 30-42 Windshield Wipers System.........................................30-78 General Description .......................................................30-78 Component Location .....................................................30-80 Wiper Motor ...............................................................30-80 Wiper Arm ..................................................................30-80 Wiper Blade ...............................................................30-80 Controls and Indications .................................................30-82
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-ii
CS130_M5_30_IceRainTOC.fm
30 - Ice and Rain Protection
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7:
Ice Detection System ............................................30-3 Ice Detector ...........................................................30-5 Ice Detection System Schematic...........................30-7 Ice Detection System Test ..................................30-11 Wing Anti-Ice System ..........................................30-13 Wing Anti-Ice Air Supply......................................30-15 Wing Anti-Ice System Leading Edge Components.................................30-17 Figure 8: Integrated Air System Controllers .......................30-19 Figure 9: Wing Anti-Ice Controls.........................................30-21 Figure 10: Wing Anti-Ice Indications.....................................30-23 Figure 11: Integrated Air System Controller .........................30-25 Figure 12: Wing Anti-Ice System Schematic ........................30-27 Figure 13: Wing Anti-Ice Activation Advanced Logic............30-29 Figure 14: Wing Anti-Ice System Test ..................................30-31 Figure 15: Wing Anti-Ice Valve .............................................30-37 Figure 16: Cowl Anti-Ice System ..........................................30-39 Figure 17: Cowl Anti-Ice System Component Location ........30-41 Figure 18: Swirl Nozzle.........................................................30-43 Figure 19: Cowl Anti-Ice System Controls............................30-45 Figure 20: Cowl Anti-Ice System Indications ........................30-47 Figure 21: Cowl Anti-Ice System Detailed Description .........30-49 Figure 22: Cowl Anti-Ice Valves ...........................................30-53 Figure 23: Probe Heat System .............................................30-55 Figure 24: Probe Heat System Components........................30-57 Figure 25: Probe Heat Control..............................................30-59
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 26: Probe Heat Electrical Interface Detailed Description............................................ 30-61 Figure 27: Windshield and Window Heat System................ 30-69 Figure 28: Windshield, Side Window, and Windshield Ice Protection Controllers................. 30-71 Figure 29: Window Heat Panel ............................................ 30-73 Figure 30: Windshield and Window Heat System Schematic ..................................... 30-75 Figure 31: Windshield Wiper System................................... 30-79 Figure 32: Windshield Wiper Motor, Wiper Arm, and Wiper Blade ................................................. 30-81 Figure 33: Windshield Wipers System Controls................... 30-83
TECHNICAL TRAINING MANUAL
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30-iii
CS130_M5_30_IceRainLOF.fm
ICE AND RAIN PROTECTION - CHAPTER BREAKDOWN ICE AND RAIN PROTECTION CHAPTER BREAKDOWN
Windshield and Window Heat System
Ice Detection System
1
5
Windshield Wiper System
Wing Anti-Ice System
2
Cowl Anti-Ice System
3
Probe Heat System
4
6
30 - Ice and Rain Protection 30-81 Ice Detection System
30-81 ICE DETECTION SYSTEM CAUTION GENERAL DESCRIPTION The detection system detects the supercooled liquid water conditions present in the clouds that can cause ice formation on the aircraft.
As a safety precaution do not touch the probes with bare hands, as they could be hot and inflict burns.
The ice detectors provide independent measurements of icing conditions. The system is comprised of two magnetostrictive ice detectors. The ice detector is a one-piece, vibrating sensing element (probe) type device. When ice accumulates on the probe, the drop in frequency triggers the ice detector to generate discrete signals (ice). In addition, heaters inside the probe and strut automatically come on to melt the ice, then turn off. With the heaters off, ice is again allowed to build up. The cycle of accumulation and melting is repeated until ice is no longer present. The discrete signals are switched off 2 minutes later. The detectors operate using 115 VAC electrical power and are always powered when the aircraft is powered. The left ice detector is powered by the AC BUS 1 and the right is powered by the AC BUS 2. When ice is detected by either detector, discrete signals are automatically sent via the data concentrator unit module cabinet (DMC) to the wing anti-ice (WAIS) system and the cowl anti-ice system (CAIS) to provide anti-icing. It also signals the three primary flight control computers (PFCCs) to modify the stall warning protection. As a backup the ice detection signal is also sent to the WAIS through the control and distribution cabinets (CDCs). The ice signals and ice detector status information are sent to the engine indication and crew alerting system (EICAS) and the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-2
CS130_30_L2_IDS_General.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
LEFT ICE DETECTOR
RIGHT ICE DETECTOR
LEFT and RIGHT CAIS
Test
AC BUS 1
Test
AC BUS 2
EICAS Ice Detector OK
Ice Detector OK DMC 1
OMS
Ice Detected 1 and Ice Detected 2
DMC 2 Ice Detected 1 and Ice Detected 2
PFCCs
WAIS
Ice Detected 3
CDCs
Ice Detected 3
LEGEND Cowl Anti-Ice System Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Primary Flight Control Computer Wing Anti-Ice System ARINC 429 Discrete
CS1_CS3_3081_013
CAIS CDC DMC PFCC WAIS
Figure 1: Ice Detection System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-3
CS130_30_L2_IDS_General.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
COMPONENT LOCATION There are two identical ice detectors installed in the system. ICE DETECTOR The ice detectors are located on each side of the forward fuselage, one is on the left, and the other on the right.
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30-4
CS130_30_L2_IDS_ComponentLocation.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
Mounting Flange
Strut
Probe
Air Flow Direction
NOTE Left ice detector shown. Right ice detector similar.
CS1_CS3_3080_002
ICE DETECTOR
Figure 2: Ice Detector Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-5
CS130_30_L2_IDS_ComponentLocation.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
DETAILED DESCRIPTION The ice detector is a one-piece, vibrating sensing element (probe) type device. An internal electrical oscillator circuit drives the sensing element to vibrate at its mechanical resonance (tuned to 40,000 Hz) using the principles of magnetostriction. The shift in the measured frequency, which corresponds to an ice thickness of approximately 0.51 mm (0.02 in.) triggers the ICE DETECTED discrete. When ice is detected, the strut and probe are de-iced and melt ice accumulation. The aerodynamic airflow then sheds the ice and provides the probe cooling for continued sensing. When it is powered up and operational, the ice detector provides an ICE DETECTOR OK signal. When icing is detected, each detector provides three signals: ICE DETECTED 1, ICE DETECTED 2, and ICE DETECTED 3. The ICE DETECTED 1 and ICE DETECTED 2 signals are sent to the DMCs and distributed to the left and right electronic engine controls (EECs) for cowl anti-ice (CAI) and to the three primary flight computers (PFCCs). The DMC also sends the ice detection signals to the integrated air system controllers (IASCs) to enable wing anti-ice system (WAIS) activation. ICE DETECTED 3 signal is sent through the control and distribution cabinets (CDCs) for the WAIS. The ICE DETECTED and the ICE DETECTOR OK signals are used to generate caution and advisory EICAS messages. The ice detection system is also monitored by the OMS. An ICE caution message is displayed when ice is detected and the wing anti-ice system (WAIS) or cowl anti-ice system (CAIS) are inoperative. An ICE advisory message indicates normal operation of wing and cowl anti-ice systems in icing conditions. When the ICE advisory message is removed the crew can select the anti-ice systems off. If an ice detector fails, the associated ICE DET FAIL caution message is displayed. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-6
CS130_30_L3_IDS_DetailedDescription.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
CDC 3-14-5
CDC 4-13-4 L ICE DET
AC BUS 1
R ICE DET AC BUS 2
SSPC 5A
Logic: Always On
SSPC 5A
Logic: Always On
LEFT ICE DETECTOR
L CAIS
R CAIS
Test
Test L EEC
R EEC
Ice Detector OK
Ice Detector OK
EICAS OMS
Ice Detected 1 Ice Detected 2
RIGHT ICE DETECTOR
DMC 1
Ice Detected 1 DMC 2
Ice Detected 2
PFCCs
CAIS CDC DMC EEC IASC PFCC WAIS
Cowl Anti-Ice System Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Engine Control Integrated Air System Controller Primary Flight Control Computer Wing Anti-Ice System AFDX ARINC 429 Discrete TTP BUS
IASC 1
Ice Detected 3
WAIS
IASC 2
CDC 1
CDC 2
CDC 3
CDC 4
Ice Detected 3
CS1_CS3_3081_014
LEGEND
Figure 3: Ice Detection System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-7
CS130_30_L3_IDS_DetailedDescription.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the ice detection system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-8
CS130_30_L2_IDS_MonitoringTests.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L ICE DET FAIL
Left ice detector has failed.
R ICE DET FAIL
Right ice detector has failed.
ICE
Ice detected and wing anti-ice or cowl anti-ice are failed or is off.
Table 2: ADVISORY Messages MESSAGE ICE
LOGIC Ice detected but ice detection, wing anti-ice and cowl anti-ice are working properly.
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30-9
CS130_30_L2_IDS_MonitoringTests.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
INITIATED BUILT-IN TEST Once powered, each ice detector initiates its own power-up built-in test (PBIT) lasting about 7 seconds. The detector produces an ICE DETECTOR OK signal after internal testing is successful and remains active, provided it continues to function normally. An initiated built-in test (IBIT) function performs a diagnostic check of both ice detectors, including communication with interconnecting systems. The IBIT is initiated from the EICAS AVIONICS page on the AVIO tab. When IBIT is initiated: •
A message IN PROG is provided
•
The ice detectors perform their internal tests
•
The data concentrator unit module cabinets (DMCs) monitor ice detection signals, ICE DETECTED 1, and ICE DETECTED 2, from the ice detectors
•
The DMCs monitor ice detection signal, ICE DETECTED 3, from the control and distribution cabinets (CDCs)
When the ice detectors internal test is complete, the detector produces an ICE DETECTOR OK signal if internal testing is successful. One of two messages, PASS or FAIL, is displayed on the AVIONICS synoptic page on the AVIO tab and more data is available in the onboard maintenance system (OMS) The PASS or FAIL message is triggered by the value of ICE DETECTOR OK signal. At any time, if an ice detector has an internal fault, the ice detector turns off and a L (R) ICE DET FAIL caution message is displayed.
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30-10
CS130_30_L2_IDS_InitiatedBIT.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
SYNOPTIC PAGE - AVIONIC
LEGEND CDC DMC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet ARINC 429 Discrete
STATUS
AIR IR
FUEL FUE EL
ELEC C
HYD YD
INFO FO
CBB
CTP CT TP CTP
AVIO
TEST STATUS
FLT CTRL CTRL
SMS RUNWAY
ENABLED
Status
Condition
PASS
Test successfully completed
FAULT
Test failure with fault message
FAIL
Test failure
IN PROG
Test in progress
----
Test invalid or aborted
VSPEEDS...
TEST
LEFT ICE DETECTOR
AURAL AURA URA AL
WING ING A/IC A/I A/ICE CE
LAMP P
PASS ICE DETECT
TCAS CAS
RIGHT ICE DETECTOR
FIRE FIR E
Test
Test
WXR
FLT CTRL CTRL
TAWS S
SHAKER R
Ice Detected 1 and Detected Ice 2
Ice Detector OK DMC 1
Ice Detected 3
Test
Test
CDCs
DMC 2
Ice Detected 1 and Detected Ice 2
Ice Detected 3
CS1_CS3_3081_015
Ice Detector OK
Figure 4: Ice Detection System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-11
CS130_30_L2_IDS_InitiatedBIT.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
30-12 WING ANTI-ICE SYSTEM GENERAL DESCRIPTION The wing is protected from icing conditions by the wing anti-ice system (WAIS), which prevents the formation of ice on the leading edge of wing slats no. 2, no. 3, and no. 4 by supplying the hot bleed air from the engines. Ice protection is not provided for the slats that are located inboard of the engines. The wing anti-ice valves (WAIVs) regulate the pressure to 27 psi in the associated wing anti-ice ducting. Flow is controlled by venturis. After heating the inside of the slat, the bleed air is expelled overboard underneath the wing. The WAIS includes two independent anti-icing systems for the left and right wings. Each wing system is connected to the other by a normally closed cross-bleed valve (CBV). The engine bleed air system supplies the hot air at required pressure and temperature to the distribution ducts installed in the slats. If one engine or bleed air supply fails, the bleed air system opens the CBV to supply air to the opposite wing. The integrated air system controllers (IASCs) send a wing anti-ice (WAI) enable signal to the data concentrator unit module cabinets (DMCs) to activate or deactivate the wing anti-ice system (WAIS), based on the ice detection system and WING switch position.
When ANTI-ICE WING switch is set to ON, the WAIS is activated regardless of icing conditions if the airspeed is above 60 kt on the ground or anytime in flight. The WAIS can be tested on the ground if the switch is in the ON position and the airspeed is below 60 kt. The test lasts 20 seconds. Each wing is continuously monitored for temperature and pressure by both IASCs. The wing anti-ice temperature sensor (WAITS) provides monitoring of the temperature of the bleed air supplied to the wing ducting. The pressure sensor measures system pressure from the outboard end of slat no. 4. Pressure is monitored to detect excessive high pressure that might lead to ducting failures, as well as low pressure which may indicate inadequate hot air for effective anti-icing or to detect a duct burst or leak. In case a leak is detected in the ducting network, the IASC shuts down the WAIS automatically. IASC 1 is powered by DC ESS BUS 1 and DC ESS BUS 2. IASC 2 is powered by DC ESS BUS 2 and DC ESS BUS 3. The WAIVs are powered by DC ESS BUS 2 provided by the control and distribution cabinets (CDCs). The pressure sensors are powered from DC ESS BUS 3.
The WAIS is monitored and controlled by the IASCs. With the ANTI-ICE WING switch in the AUTO position, the IASCs automatically activate the WAIS on receiving ice signals from one or both ice detectors. The WAIS is monitored and controlled by the IASCs. In AUTO mode, the WAIS is inhibited on the ground and during takeoff. In flight, the IASCs automatically activate the WAIS on receiving ice signals from one or both ice detectors.
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30-12
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
ICE DETECTION SYSTEM
NOTE IASC 1 is powered by DC ESS BUS 1 and DC ESS BUS 2. IASC 2 is powered by DC ESS BUS 2 and DC ESS BUS 3.
CDCs (DC ESS BUS 2) ANTI-ICE WING
EICAS
AUTO
OFF
ON
DMCs
SYN OMS
ANTI-ICE PANEL IASCs
LEGEND
PS SOL
Pressure Sensor
CBV SOL
SOL
Slats
Slats WAIV
WAIV
WAITS
WAITS
PS
Venturi
Solenoid Temperature Sensor
DC ESS BUS 3
PS
CS1_CS3_3012_019
CBV CDC DMC IASC WAITS WAIV
Bleed Air Wing Anti-Ice Air Cross-Bleed Valve Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve ARINC 429 Discrete
Figure 5: Wing Anti-Ice System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-13
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE DUCTING The wing anti-ice ducting consists of a venturi duct and three piccolo tubes connected by interslat hoses. The air exits the piccolo tubes and circulates around the front bay inner chamber of the slat to be expelled overboard through slots and exhaust holes in the slat rear bay. A telescopic duct connects the wing anti-ice system air duct to the piccolo tubes to allow for the slats extension and retraction.
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30-14
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Engine Bleed Air Supply
Telescopic Duct Interslat Hoses
Venturi
SOL
WAIV
WAITS
Rear Bay Seal
Front Bay
PS
Piccolo Tubes
Exhaust Holes
Piccolo Tube
WAITS WAIV
Bleed Air Wing Anti-Ice Air Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve
PS
Pressure Sensor
SOL
Solenoid
Slat
CROSS-SECTION VIEW
CS1_CS3_3012_012
LEGEND
Figure 6: Wing Anti-Ice Air Supply Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-15
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
COMPONENT LOCATION
PRESSURE SENSORS
The wing anti-ice system (WAIS) consists of the following components:
Each pressure sensor is connected by a tube, to the outboard end of the piccolo tube assembly in slat no. 4.
•
Wing anti-ice valves (WAIVs)
•
Venturis
•
Telescopic ducts
•
Piccolo tubes
•
Wing anti-ice temperature sensors (WAITSs)
•
Pressure sensors (PSs)
•
Integrated air system controllers (IASCs) (Refer to figure 8).
WING ANTI-ICE VALVES Each wing anti-ice valve (WAIV) is located in the wing leading edge, outboard of the engine pylon, forward of the front spar. VENTURIS A venturi is located immediately downstream and outboard of the WAIV, in the wing leading edge cavity aft of slat no. 2. TELESCOPIC DUCTS Each telescopic duct is located in the wing leading edge cavity aft of slat no. 2. PICCOLO TUBES On each wing, slat no. 2, no. 3, and no. 4 each contain a piccolo tube. The piccolo tubes are joined by interslat hoses. WING ANTI-ICE TEMPERATURE SENSORS Each wing anti-ice temperature sensor (WAITS) is installed downstream of the venturi duct, before the telescopic duct.
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30-16
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
B
A Slat No. 2
Wing Anti-Ice Valves (2x)
A Slat No. 3
C Slat No. 4
Venturis (2x)
Telescopic Ducts (2x)
Wing Anti-Ice Temperature Sensor
A Piccolo Tubes (6x)
Interslat Hose
Pressure Sensors (2x)
NOTE Left wing shown. Right wing similar. Slats not shown for clarity.
C
CS1_CS3_3012_013
B
Figure 7: Wing Anti-Ice System Leading Edge Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-17
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
INTEGRATED AIR SYSTEM CONTROLLERS There are two integrated air system controllers (IASCs) located in the mid equipment bay.
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30-18
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
IASC 2
IASC 1
INTEGRATED AIR SYSTEM CONTROLLERS (2X)
CS1_CS3_3012_017
MID EQUIPMENT BAY FWD RACK
Figure 8: Integrated Air System Controllers Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-19
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
CONTROLS AND INDICATIONS ANTI-ICE CONTROL PANEL The wing anti-ice system (WAIS) has a three-position rotary switch located on the overhead panel of the flight deck. The switch is normally left in the AUTO position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-20
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
A
OVERHEAD PANEL
ANTI-ICE L COWL
WING
AUTO
OFF
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
CS1_CS3_3012_021
ANTI-ICE PANEL
Figure 9: Wing Anti-Ice Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-21
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE INDICATIONS The wing anti-ice system (WAIS) indications and messages are displayed on the EICAS page and the AIR synoptic page. EICAS Page The wing anti-ice (WAI) icon is displayed on EICAS when the system is not OFF. The WAI icon is displayed in green when the system is activated, displayed in white when inhibited, or amber when a failure is reported. AIR Synoptic Page The status of the wing anti-ice valves (WAIVs) and the wing leading edge are displayed on the AIR synoptic page.
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30-22
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
CLB
STATUS TUS
92.2
92.2
73.0
73.3
FUEL FUE EL
722
EGT N2 FF (KPH) OIL TEMP OIL PRESS
ELECC
WING ANTI-ICE VALVE
FLT CTR CTRLL
HYD YD
AVIONIC NIC
INFO
CB
Condition Open with flow Closed
718
86.5 1090 124 116
DOOR OR
Symbol
N1
CAI WAI
AIR IR
86.6 1110 124 116
LO
FWD 22 °C 22 °C -+ 2
AFT 22 °C 22 °C -+ 2
10 °C °
15 °CC
15 °CC
CKPT 23 °C 22 °C
CARGO
15°CC
Open with no flow (normal) Open with no flow (fail) Open with flow (fail)
CAI WAI
RAM AIR
TRIM AIR L PACK
HI
R PACK
Invalid Closed (fail)
HI
EICAS WING ANTI-ICE STATE
45
WING ANTI-ICE SYSTEM Symbol
Condition
WAI
Wing anti-ice activated
WAI
Wing anti-ice inhibited
WAI
Wing anti-ice caution failure or overheat
WAI
WAI failed in warning condition
PSI
XBLEED
45 PSI
Symbol
Condition Off Normal
Left Wing Anti-Ice Valve
Abnormal
Right Wing Anti-Ice Valve
SYNOPTIC PAGE – AIR
CS1_CS3_3012_001
APU
Figure 10: Wing Anti-Ice Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-23
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
DETAILED DESCRIPTION INTEGRATED AIR SYSTEM CONTROLLERS The integrated air system controllers (IASCs) control and monitor wing anti-ice system (WAIS) operation using position sensing of the wing ant-ice valve (WAIV) (full closed), wing anti-ice (WAI) temperature and pressure. Each IASC has three channels: channel A, channel B, and a safety channel. Channel A is not used by the WAIS. The WAIS uses the IASC to: •
Inhibit or enable activation
•
Monitor WAIS pressure and temperature
•
Monitor WAIS component health
When the aircraft is powered, the active IASC is determined from odd dates for IASC 1 and even dates for IASC 2. Channel B of the active IASC for the day is assigned to WAIS activation. The safety channels of both IASCs, in conjunction with channel B of the standby IASC, monitor WAIS temperature and pressure and provide signals to the EICAS, through the data concentrator unit module cabinets (DMCs). The standby IASC also monitors WAIS component health by performing a continuous built-in test (CBIT). Data from CBIT is reported to the onboard maintenance system (OMS) through the DMCs The operation of the WAIS is shown schematically on the AIR synoptic page. Synoptic page data comes from the safety channel through the DMCs.
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CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Channel A
Channel B
Safety Channel
L/R WAIS Activation
L/R WAIS Monitoring: • Pressure • Temperature L/R Component Health Monitoring
IASC (Standby) L PRESSURE SENSOR
L WAITS
L WAIV
EICAS DMC SYN
R PRESSURE SENSOR
R WAITS
OMS
R WAIV
L/R WAIS Monitoring: • Pressure • Temperature L/R Component Health Monitoring
Channel A
Channel B
Safety Channel
IASC (Active)
LEGEND DMC IASC OMS WAIS WAITS WAIV
Data Concentrator Unit Module Cabinet Integrated Air System Controller Onboard Maintenance System Wing Anti-Ice System Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve ARINC 429
CS1_CS3_3012_010
L/R WAIS Activation
Figure 11: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-25
CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE OPERATION The wing anti-ice system (WAIS) uses precooled engine bleed air, sprayed by the piccolo tubes onto the internal surface of the wing leading edge slats to prevent ice buildup. The bleed air system controls engine bleed air supply pressure and temperature to the WAIS. The supply bleed air temperature target at the bleed temperature sensor (BTS) is nominally 227°C (440°F) in flight. On the ground, during initiated built-in test (IBIT), the temperature target is 200°C (392°F). The WAIV controls the pressure of bleed air to the wing leading edge slats. Wing anti-ice temperature sensors (WAITSs) are used by the integrated air system controller (IASC) to monitor the bleed air temperature, downstream of the venturis. Downstream of the wing anti-ice valve (WAIV), the hot bleed air passes through the venturi and telescopic duct to reach the slat no. 2 piccolo tube, then to slat no. 3 and no. 4 through interslat hoses. The area between the slats is not heated.
The pressure sensor detects pressure at the end of the piccolo tubes of slats no. 4. A low-pressure condition of less than 14 psi for more than 15 seconds generates a WING ANTI-ICE FAIL warning message on EICAS for the associated wing. If the low-pressure condition persists and results in a low-temperature condition on the leading edge, the WING ANTI-ICE LO HEAT caution message replaces the previous warning message. If an overpressure condition of more than 21 psi for more than 15 seconds (indicative of a WAIV failed open condition) is detected by the pressure sensor (PS), a WING ANTI-ICE FAIL warning message is displayed on EICAS for the associated wing. This message also appears if the WAIV if not closed when on the ground. If the WAIV position does not indicate full closed with the WAIS off, a WING ANTI-ICE FAIL ON caution message is displayed on EICAS for the associated wing.
The purpose of the venturi is to choke the air flow in case of a burst or crack in the downstream ducting. The dual-channel WAITS monitors temperature downstream of the WAIV. If a high-temperature condition is detected that exceeds 254°C (489°F) for more than 30 seconds, a WING ANTI-ICE OVHT caution message is displayed on EICAS for the associated wing. If a low-temperature condition of less than 200°C (392°F) is detected by the WAITS for more than 60 seconds, a WING ANTI-ICE LO HEAT caution message is displayed on EICAS for the associated wing. The actual low-temperature threshold is a function of aircraft altitude and pressure reported by the pressure sensor.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
ICE DETECTION SYSTEM
1 CDC 2-5-17 WING A-ICE VLV DC ESS SSPC 3A BUS 2 Logic: Combinational CDC 2
Ice Detected 3
CDC 3 and CDC 4
CDC 2
1 ANTI-ICE WING
Ice Detected 1
OFF
Ice Detected 2
EICAS
AUTO
ON
DMCs
SYN ANTI-ICE PANEL OMS
IASC 1
IASC 2
LEGEND
PS SOL
Pressure Sensor
Solenoid
CBV
SOL
Slats
SOL
Slats
WAIV
WAIV
WAITS
WAITS BTS CB-D2 PS
DC ESS BUS 3
Venturi ECS SAFETY SNSR
Temperature Sensor
PS
CS1_CS3_3012_011
BTS CBV CDC DMC IASC WAITS WAIV
Bleed Air Wing Anti-Ice Air Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve AFDX ARINC 429 Discrete TTP BUS
Figure 12: Wing Anti-Ice System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-27
CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE ACTIVATION CONTROL The WAIVs are commanded to simultaneously open when they receive 28 VDC from control and distribution cabinet 2 (CDC 2) and controlled by advanced logic. The combinational logic within CDC 2 receives input from the following sources: •
An enable/inhibit signal from the IASCs via the DMCs
•
The WING anti-ice switch
•
Ice detection signals from the ice detection system
•
Airspeed < 60 kt and WAIS test selected
-
Airspeed > 60 kt Anytime the switch is in ON position
With WAIS switch in AUTO position, and ice is detected, the system turns ON: •
Main engine start (MES)
•
Bleed air leak detected by bleed air leak and overheat detection system (BALODS)
•
Wing anti-ice (WAI) switch selected OFF
When in flight, the WAIS can be safely operated below 12°C (53°F) total air temperature (TAT) or airspeed above 60 kt. A WING A/ICE ON caution message is provided by the system when the system is ON and the TAT is above 15°C (59°F).
On ground: -
•
•
Any loss of redundant or non-critical function results in a WING A/ICE FAULT advisory message.
In flight: -
No engine bleed is available (one engine minimum required)
When the WAIS rotary switch is set to ON or AUTO and neither engine is running, a WING A/ICE MISCONFIG caution message is displayed on EICAS.
On ground: -
•
The WAIVs are closed 2 minutes after the ICE DETECT signal is no longer present.
With the WAIS switch in the ON position, the system turns ON: •
The WAIS is automatically deactivated when any of the following occurs:
Never
In flight, in either of the following conditions: -
Altitude > departure airport altitude + 457 m (1,500 ft)
-
Takeoff and 2 minutes after airspeed > 60 kt
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
AUTO MODE
MANUAL MODE
Engine Bleed Air ACFT on Ground Airspeed < 60 kt WAIS Test Enabled WAIS Switch ON
Takeoff Airspeed > 60 kt + 2 Minutes
20 Second Timer
ALT. > Departure Airport ALT. + 1.500 ft WAIS Valve Open
WAIS Valve Open
Aircraft in Flight Ice Detected WAIS Switch in AUTO
CS1_CS3_3012_020
ACFT in Flight
LEGEND WAIS
Wing Anti-Ice Switch
Figure 13: Wing Anti-Ice Activation Advanced Logic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-29
CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
MONITORING AND TESTS SYSTEM TEST A initiated built-in test (IBIT) is available on the AVIONICS synoptic page on the AVIO tab. The test can be carried out on ground, with the bleed air system pressurized by the engines. IBIT is limited to one cycle prior to takeoff to protect the leading edges from damage caused by excessive heat cycles. The airplane flight manual (AFM) limits the test to once per day. In case of test failure, the AFM limits the time between tests to a minimum of 30 minutes. The WING anti-ice switch must be set to the ON position for the test. If the auxiliary power unit (APU) bleed air is being used, the IASCs automatically switch to engine bleed. The icon is grayed out if conditions are not correct for the IBIT. When the WING A/ICE icon is pressed, the IASCs open the WAIVs and monitor the pressures obtained. The test period is limited to 20 seconds, after which the WAIVs close. The initiated built-in test (IBIT) results are displayed on the synoptic page as either PASS, FAIL, or FAULT. IBIT does not activate if a failure is detected and not rectified, indicated by the FAIL message. After failure rectification, the test can be re-initiated by resetting the IASCs. A wait time of 30 minutes is required to ensure cooling of the wing slats before repeating the test. A display of ----- indicates the system had communication problems during the test. Additionally, the IASCs conduct a power-up built-in test (PBIT) when power is first available on the aircraft. This check determines the validity of the two pressure sensor signals.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-30
CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
TEST STATUS Status
Condition
PASS
Test successfully completed
FAULT
Test failure with fault message
FAIL
Test failure
IN PROG
Test in progress
----
Test invalid or aborted
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
CTP
AVIO SMS RUNWAY
ENABLED VSPEEDS...
Engines Running On Ground
TEST
AURAL
PASS WING A/ICE
LAMP
ICE DETECT
TCAS
FIRE
WXR
FLT CTRL
TAWS
SHAKER
L COWL
WING
AUTO
OFF
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
CS1_CS3_3012_005
ANTI-ICE
Figure 14: Wing Anti-Ice System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-31
CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
CAS MESSAGES The following page provides the crew alerting system (CAS) and INFO messages associated with the wing anti-ice system (WAIS).
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Table 6: STATUS Messages Table 3: WARNING Messages MESSAGE
MESSAGE
LOGIC
L WING A/ICE FAIL
L WAI not closed in any condition on ground out of test sequence, overpressure or low-pressure.
R WING A/ICE FAIL
R WAI not closed in any condition on ground out of test sequence, overpressure or low-pressure.
WING A/ICE ON
Wing anti-ice selected ON.
WING A/ICE OFF
Wing anti-ice selected OFF.
Table 7: INFO Messages MESSAGE 30 WING A/ICE FAULTWING A/ICE TEMP SNSR REDUND LOSS
Table 4: CAUTION Messages MESSAGE
LOGIC
LOGIC
LOGIC Single failure of one WAITS channel.
L WING A/ICE OVHT
Left wing bleed overtemperature detected.
R WING A/ICE OVHT
Right wing bleed overtemperature detected.
30 L WING A/ICE LO HEAT L WAITS drift low. - L WING A/ICE TEMP SNSR INOP
L WING A/ICE LO HEAT
Low combination of pressure and temperature for L WAI (also function of altitude).
30 L WING A/ICE LO HEAT L bleed temperature control too low for WAI. - CTRL TEMP INOP
R WING A/ICE LO HEAT
Low combination of pressure and temperature for R WAI (also function of altitude).
30 L WING A/ICE LO HEAT L HPV failed closed leading to WAI low - L HPV FAIL CLSD temperature.
WING A/ICE FAIL
Left or right WAI function failed.
WING A/ICE ON
Wing anti-ice selected ON and TAT above 15°C (59° F).
30 L WING A/ICE OVHT - L L WAITS drift high. WING A/ICE TEMP SNSR INOP
WING A/ICE MISCONFIG
WAI selected ON or WAI AUTO and ice detected with both engine bleeds OFF.
30 L WING A/ICE OVHT CTRL TEMP INOP
L bleed temperature control too high for WAI.
30 R WING A/ICE LO HEAT - R WING A/ICE TEMP SNSR INOP
R WAITS drift low.
Table 5: ADVISORY Messages MESSAGE WING A/ICE FAULT
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC Loss of redundant or non-critical function for the wing anti-ice system or WAIV high leakage. Refer to INFO messages.
30 R WING A/ICE LO R bleed temperature control too low for WAI. HEAT - CTRL TEMP INOP 30 R WING A/ICE LO HEAT - R HPV FAIL CLSD
R HPV failed closed leading to WAI low temperature.
30 R WING A/ICE OVHT R WING A/ICE TEMP SNSR INOP
R WAITS drift high.
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Page Intentionally Left Blank
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Table 7: INFO Messages MESSAGE
LOGIC
30 R WING A/ICE LO R bleed temperature control too low for WAI. HEAT - CTRL TEMP INOP 30 WING A/ICE FAULT Loss of ICE DETECTED signal at IASC input. WING A/ICE AUTO MODE INOP 30 WING A/ICE FAULT - L Erroneous L WAIV switch acquisition on one WING A/ICE VLV POS SW channel. INOP 30 WING A/ICE FAULT - R Erroneous R WAIV switch acquisition on one WING A/ICE VLV POS SW channel. INOP 30 WING A/ICE FAIL - L WING PRESS FAIL
Left wing anti-ice overpressure or low-pressure.
30 WING A/ICE FAIL - R WING PRESS FAIL
Right wing anti-ice overpressure or low-pressure.
30 WING A/ICE FAULT - L WING A/ICE VLV LEAK
Left WAIV internal leakage.
30 R WING A/ICE FAIL - R Right WAIV internal leakage. WING A/ICE VLV LEAK 30 L WING A/ICE FAIL - L WING A/ICE VLV FAIL OPEN
Left wing anti-ice failed open.
30 R WING A/ICE FAIL - R Right wing anti-ice failed open. WING A/ICE VLV FAIL OPEN 30 WING A/ICE FAULT - L WING A/ICE PRESS SNSR INOP
Left outboard pressure sensor out of range or drifted.
30 R WING A/ICE FAIL - R Right outboard pressure sensor out of range or WING A/ICE VLV LEAK drifted.
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
PRACTICAL ASPECTS WING ANTI-ICE VALVE For dispatch purposes, the wing anti-ice valve (WAIV) is equipped with a manual override provision to rotate the valve to the closed position. A deactivation screw locks the valve in the closed position. When the deactivation screw is removed and used to secure the valve closed, air is bled from the pneumatic actuator line to prevent the valve from operating.
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30-36
CS130_30_L2_WAIS_PracticalAspects.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Deactivation Screw in Stowed Position
Deactivation Screw Position for Valve Lockout
CS1_CS3_3011_003
Manual Override
Figure 15: Wing Anti-Ice Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-37
CS130_30_L2_WAIS_PracticalAspects.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
30-22 COWL ANTI-ICE SYSTEM GENERAL DESCRIPTION The cowl anti-ice system (CAIS) uses engine bleed air from the 6th stage compressor. The air is supplied to the engine inlet cowl, heating the leading edge and preventing ice accumulation.
The EEC activates the system based on switch settings from the ANTI-ICE panel and the ice detection system signals. The CAIS operates when the EEC commands a pair of CAIVs to open.
The cowl anti-ice valves (CAIVs) provide a pressure regulating and shutoff function. The two valves are similar in design, and provide redundancy in case of the failure of a single valve, however, the CAIVs are not interchangeable and have different part numbers.
A dual-element fan cowl temperature sensor in the fan cowl sends fan case area temperature to the EEC. This sensor is used to detect a temperature rise due to a possible CAIS duct rupture.
Each CAIV is a spring-loaded, failsafe open, poppet-type valve. Closure of the valve requires electrical power applied to the solenoid assembly and bleed air pressure. The valves are open when the solenoid assembly is electrically de-energized and/or there is a loss of bleed air pressure. Each engine CAIS operates independently. The COWL rotary switches located on the ANTI-ICE panel, activates the CAIS when set to the AUTO position, or when selected to the ON position in manual mode. Primary power for both electronic engine controls (EECs) is from their respective engine-driven permanent magnet generator (PMG) supplying 3-phase 34 VAC. Each EEC channel automatically switches power source to aircraft power source should the PMG fail. The left EEC is powered by DC ESS BUS 3 for channel A, and DC ESS BUS 1 for channel B. The right EEC is powered by DC ESS BUS 3 for channel A, and DC ESS BUS 2 for channel B.
A dual-channel pressure sensor sends CAIS duct pressure to the EEC for monitoring CAIS operation. When the engine reaches ground idle after starting, the EEC commands one of the CAIVs to close and monitors downstream bleed pressure, confirming its ON/OFF capability. Each CAIV is checked alternately every other flight. System messages are sent by the EEC to the engine indication and crew alerting system (EICAS), and the onboard maintenance system (OMS) through the data concentrator unit module cabinets (DMCs). System operation is displayed schematically on the AIR synoptic page.
The operational logic for the CAIS is contained in the EEC, operating on two channels, A and B. Further details on the EEC can be found in ATA 73 Engine Fuel and Control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_30_L2_CAIS_General.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
ANTI-ICE L COWL
WING
AUTO
OFF
Fan Cowl Temperature Sensor
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
Fan Cowl Temperature Sensor
ANTI-ICE PANEL EICAS
Pressure Sensor
Pressure Sensor
SYN OMS 6th Stage Air
PMG
PMG PS
PS
DMCs
EEC
EEC
Ch B
Ch A
ICE DETECTION SYSTEM
EEC
EEC
Ch A
Ch B
6th Stage Air
Cowl Anti-Ice Valve (2x)
Cowl Anti-Ice Valve (2x) DC ESS BUS 2 DC ESS BUS 3 DC ESS BUS 1
Bleed Air Data Concentrator Unit Module Cabinet Electronic Engine Control Permanent Magnet Generator (3-Phase 34 VAC) ARINC 429 Discrete
DC ESS BUS 3 CS1_CS3_3021_006
DMC EEC PMG
LEGEND
Figure 16: Cowl Anti-Ice System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-39
CS130_30_L2_CAIS_General.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
COMPONENT LOCATION The cowl anti-ice system (CAIS) consists of left and right subsystems each of which contain: •
Cowl anti-ice valves (CAIVs)
•
Pressure sensor
•
Fan cowl temperature sensor
•
Swirl nozzle
COWL ANTI-ICE VALVES The two cowl anti-ice valves (CAIVs) are mounted one above the other at the 6 o’clock position under the engine compressor case. PRESSURE SENSOR The pressure sensor is located on the engine fan case aft section, at the 5 o’clock position. FAN COWL TEMPERATURE SENSOR The fan cowl temperature sensor is located on the right side of the engine fan case at the 4:30 o’clock position. SWIRL NOZZLE The swirl nozzle duct is located within the engine inlet cowl.
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CS130_30_L2_CAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
A COWL ANTI-ICE VALVES (2X)
B PRESSURE SENSOR
A B
C
C FAN COWL TEMPERATURE SENSOR
D
SWIRL NOZZLE
CS1_CS3_3021_007
D
Figure 17: Cowl Anti-Ice System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-41
CS130_30_L2_CAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
DETAILED COMPONENT INFORMATION SWIRL NOZZLE The swirl nozzle has an inner wall and an outer wall. Hot air exiting the nozzle swirls within the inlet cowl, heating the surface of the inlet cowl lip to provide anti-icing. In the event of a rupture to the inner wall, the outer wall contains the hot engine bleed air directing it backwards toward an insulation boot. The insulation boot is designed to allow hot air from a ruptured inner wall to blow back into the engine fan casing area. The rupture is detected by the fan cowl temperature sensors.
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30-42
CS130_30_L3_CAIS_DetailedComponentInfo.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
Inlet Cowl Aft Bulkhead
Insulation Boot (deforms when hot) Air Flow
Hot Air Leak
Exhaust Slots Inner Wall
SWIRL NOZZLE WITH LEAK
Outer Wall
CS1_CS3_3021_008
Fan Cowl Temperature Sensor
Figure 18: Swirl Nozzle Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-43
CS130_30_L3_CAIS_DetailedComponentInfo.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
CONTROLS AND INDICATIONS The cowl anti-ice system (CAIS) is controlled by two switches located on the anti-ice panel to provide separate controls for the left, and right systems.
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CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
A
OVERHEAD PANEL
ANTI-ICE L COWL
WING
AUTO
OFF
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
CS1_CS3_3021_009
ANTI-ICE PANEL
Figure 19: Cowl Anti-Ice System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-45
CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
INDICATIONS Indication of cowl anti-ice system (CAIS) operation is displayed on EICAS and the AIR synoptic page. The cowl anti-ice (CAI) icon is displayed on EICAS to advise the crew of the system status. The CAI icon on EICAS is displayed in GREEN when the system is activated, displayed in WHITE when inhibited, or AMBER when a failure is reported.
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30-46
CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
CLB
STATUS TUS
92.2
92.2
73.0
73.3
718
722
EGT
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
DOOR OR
ELECC
FLT CTR CTRLL
COWL ANTI-ICE VALVE Symbol
FUEL FUE EL
HYD YD
AVIONIC NIC
INFO
CB
Condition Open with flow Closed
N1
CAI WAI
AIR IR
86.6 1110 124 116
LO
FWD 22 °C 22 °C -+ 2
AFT 22 °C 22 °C -+ 2
10 °C °
15 °CC
15 °CC
CKPT 23 °C 22 °C
CARGO
15°CC
Open with no flow (normal) Open with no flow (fail) Open with flow (fail)
CAI WAI
RAM AIR
TRIM AIR L PACK
HI
R PACK
Invalid Closed (fail)
HI
EICAS COWL ANTI-ICE STATE COWL ANTI-ICE ICON Symbol
45 PSI
XBLEED
45 PSI
Condition
Symbol
Condition Off
CAI
Cowl anti-ice activated
Normal
CAI
Cowl anti-ice failure reported
Abnormal
CAI
Cowl anti-ice inhibited
SYNOPTIC PAGE – AIR
CS1_CS3_3022_001
APU
Figure 20: Cowl Anti-Ice System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-47
CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
DETAILED DESCRIPTION
TEMPERATURE AND PRESSURE SENSING
The cowl anti-ice system (CAIS) uses bleed air supplied from the 6th stage of the high-pressure compressor (HPC) to heat the engine inlet cowl lip.
The fan cowl inlet temperature sensor and the cowl anti-ice (CAI) pressure sensor on each engine send signals to EEC for monitoring.
When the CAIS is in operation, the cowl anti-ice valves (CAIVs) are commanded open allowing hot air flow to the inlet cowl. The hot bleed air swirls through the inlet cowl and is discharged overboard through the exhaust louvers. If one of the two CAIVs remains closed, a L (R) COWL A/ICE FAIL caution message on the EICAS appears. In flight, when the COWL switch is selected to ON, the CAIS is activated regardless of ice detection. If the outside air temperature (OAT) is greater than 15°C (60°F), the COWL ANTI-ICE ON caution message is displayed on EICAS for the associated engine. On the ground, the CAIS is inhibited when selected to ON if the outside air temperature (OAT) is above 15°C (60°F). In flight, when the COWL switch is selected to AUTO, the CAIS is activated when ice is detected. The CAIS is turned off when the ice detectors no longer signal ice (2 minutes after leaving icing conditions). If the ice signals are not valid, the electronic engine control (EEC) assumes icing conditions and opens the CAIVs. On the ground in AUTO mode, CAIS is inhibited with takeoff thrust commanded. The EEC maintains the takeoff inhibit condition until either the aircraft has climbed 457 m (1500 ft) above departure airport elevation, or 2 minutes have elapsed since takeoff inhibit was set.
In case of a rupture to the CAIS ducting, the affected temperature sensor sends signals to the EEC. The EEC generates an ENG NACELLE OVHT caution message in the EICAS only if the EEC is not able to close the CAIV. With the CAIV commanded closed due to overheat, the EEC continues to command it closed until the EEC is reset. Unless reset, once the CAIS system is required to provide anti-ice heat in ON or AUTO mode, the EEC generates a COWL A/ICE FAIL caution message on EICAS. While in operation, if the CAIS pressure falls below 40 psi, the pressure sensor signal causes the EEC to generate a COWL A/ICE FAIL caution message. If CAIS pressure rises above 71 psi, an ENGINE FAULT advisory message is displayed. Any time an engine is running, and the EEC does not detect valid ICE DETECTION signal, the EEC opens the CAIS valves. During an engine start, both cowl anti-ice valves (CAIVs) are open until the engine reaches ground idle. Once the engine reaches ground idle, the electronic engine control (EEC) commands one of the CAIVs closed. This enables the EEC to verify system control. Each valve is checked alternately on every start. After the valve is checked, the EEC commands both valves to close.
The EEC provides system data through the data concentrator unit module cabinets (DMCs) for EICAS messages, the AIR synoptic page and for the onboard maintenance system (OMS).
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CS130_30_L3_CAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
LEGEND CAIS DMC EEC HPC
Bleed Air Cowl Anti-Ice System Data Concentrator Unit Module Cabinet Electronic Engine Control High-Pressure Compressor ARINC 429 Discrete
ICE DETECTION SYSTEM
ANTI-ICE L COWL
EICAS
WING
AUTO
OFF
SYN
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
DMCs
OMS ANTI-ICE PANEL
DC ESS BUS 1 DC ESS BUS 3
AUTO MODE Takeoff Airspeed > 60 kt + 2 Minutes ALT. > Departure Airport ALT. + 1.500 ft
EEC
Aircraft in Flight Ice Detected
CAIS Valve Open
CAIS Switch in AUTO
Fan Cowl Temperature Sensor
Aircraft On Ground Outside Air Temperature < 15°C (60°F)
Pressure Sensor
Primary Valve Secondary Valve
MANUAL MODE
CAIS Switch ON
PMG
Aircraft in Flight NOTE Left cowl anti-ice system shown. Right cowl anti-ice system similar.
LEFT ENGINE Swirl Nozzle Location
Exhaust Slots
CAIS Valve Open CS1_CS3_3021_010
6th Stage Air
Figure 21: Cowl Anti-Ice System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-49
CS130_30_L3_CAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the cowl anti-ice system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-50
CS130_30_L2_CAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
CAS MESSAGES Table 10: STATUS Messages
Table 8: CAUTION Messages MESSAGE
LOGIC
L ENG NACELLE OVHT
R ENG NACELLE OVHT
L COWL A/ICE FAIL
MESSAGE
Left burst cowl anti-ice duct or L HPC valve failed open or buffer air shutoff valve failed open. Right burst cowl anti-ice duct or R HPC valve failed open or buffer air shutoff valve failed open. Left cowl anti-ice system failed (valves closed).
R COWL A/ICE FAIL
Right cowl anti-ice system failed (valves closed).
L COWL A/ICE FAIL ON
Left cowl anti-ice system failed (valves open).
R COWL A/ICE FAIL ON
Right cowl anti-ice system failed (valves open).
COWL A/ICE ON
L or R COWL anti-ice manually selected ON while the OAT is above 15°C (59°F).
LOGIC
L COWL A/ICE ON
Left cowl anti-ice manually selected ON.
R COWL A/ICE ON
Right cowl anti-ice manually selected ON.
L-R COWL A/ICE ON
Left and right cowl anti-ice manually selected ON.
L COWL A/ICE OFF
Left cowl anti-ice manually selected OFF.
R COWL A/ICE OFF
Right cowl anti-ice manually selected OFF.
L-R COWL A/ICE OFF
Left and right cowl anti-ice manually selected OFF.
Table 9: ADVISORY Messages MESSAGE
LOGIC
L ENGINE FAULT
Loss of redundant or non-critical function for the left engine.
R ENGINE FAULT
Right of redundant or non-critical function for right engine.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-51
CS130_30_L2_CAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
PRACTICAL ASPECTS COWL ANTI-ICE VALVES A manual override feature is provided on each valve to lock a failed valve in the open position for dispatch, which leaves the remaining valve for control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-52
CS130_30_L2_CAIS_PracticalAspects.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
Inlet
Solenoid Assembly
Outlet
Inlet
Manual Override
Solenoid Assembly
CS1_CS3_3021_004
MANUAL OVERRIDE
Outlet
Figure 22: Cowl Anti-Ice Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-53
CS130_30_L2_CAIS_PracticalAspects.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
34-10 PROBE HEAT SYSTEM A PROBE HEAT pushbutton annunciator (PBA) powers the ADSP and TAT probe heaters for 2 minutes on the ground provided the auxiliary power unit (APU) is running or external power is connected to the aircraft.
GENERAL DESCRIPTION The probe heat system prevents ice accumulation on the: •
Angle-of-attack (AOA) vanes
•
Air data system probes (ADSP)
•
Total air temperature (TAT) probes
WARNING
Each AOA vane has a vane heater and a case heater. For the left AOA vane, both heaters are powered by DC BUS 1. For the right AOA vane, both heaters are powered by DC BUS 2. These heaters are on whenever an engine is running or if the aircraft is in the air.
AS A SAFETY PRECAUTION, DO NOT TOUCH THE PROBES WITH BARE HANDS, AS THEY COULD BE HOT AND INFLICT BURNS.
The ADSP and TAT probe heaters operate in two modes: self-regulating and full. The self-regulating mode provides power to these heaters once the engines are running and the aircraft is on the ground under 55 kt. In this mode the probe regulates its own temperature. When the aircraft is in the air the heaters switch to full mode, which provides unregulated heating of the probes.
Table 11: Probe Heat Power Sources
The ADSP requires AC power for its primary and secondary power sources. Primary AC power is from AC BUS 1, AC BUS 2, and AC ESS BUS. Secondary power for ADSP 1 is AC BUS 2. For ADSP 2, secondary power is from AC BUS 1. The source of secondary power for ADSP 3 and 4 is from independent and dedicated static inverters which use power from the DC ESS BUS 3 to provide AC power. The primary heating power source for both TAT probes is the AC ESS BUS via their ADSPs. The left TAT probe heater is controlled by ADSP 3. The right TAT probe heater is controlled by ADSP 4. Secondary power is provided directly from the secondary power source from associated ADSPs.
PROBE
PRIMARY POWER
SECONDARY POWER
ADSP 1
AC BUS 1
AC BUS 2
ADSP 2
AC BUS 2
AC BUS 1
ADSP 3
AC ESS BUS
DC ESS BUS 3 via static inverter 1
ADSP 4
AC ESS BUS
DC ESS BUS 3 via static inverter 2
TAT L
ADSP 3
-----
TAT R
ADSP 4
-----
AOA L
DC BUS 1
-----
AOA R
DC BUS 2
----CS1_CS3_TB30.001
ADSP and TAT heater messages are displayed on the engine indication and crew alerting system (EICAS) and monitored by the onboard maintenance system (OMS). The AOA vane heaters are not monitored.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-54
CS130_30_L2_PHS_General.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
DC BUS 1 AC BUS 1
AC BUS 2
Primary AC Power
Secondary AC Power
Case
Vane
EICAS OMS ADSP 1
LEFT AOA VANE
PROBE HEAT AC ESS BUS
Primary AC Power
2 Minute Timer
GND ON
DC ESS BUS 3
Secondary AC Power
LEGEND ADSP AOA EEC TAT
Air Data System Probe Angle-of-Attack Electronic Engine Control Total Air Temperature
ADSP 3
NOTE Left side shown. Right side similar.
STATIC INVERTER 1
CS1_CS3_3030_003
LEFT TAT
Figure 23: Probe Heat System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-55
CS130_30_L2_PHS_General.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
COMPONENT LOCATION The probe heat system consists of the following: •
Angle-of-attack (AOA) vanes
•
Air data system probes (ADSPs)
•
Total air temperature (TAT) probes
•
Static inverters
ANGLE-OF-ATTACK VANE There are two angle-of-attack (AOA) vanes, one on each side of the forward fuselage. AIR DATA SYSTEM PROBE The are four air data system probes (ADSPs), two on each side of the forward fuselage. TOTAL AIR TEMPERATURE PROBE There are two total air temperatures (TATs) probes, one on each side of the forward fuselage. STATIC INVERTER There are two static inverters, one on each side of the forward equipment bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-56
CS130_30_L2_PHS_ComponentLocation.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
AOA Vane
TAT Probe
LEGEND ADSP AOA TAT
NOTE
Air Data System Probe Angle-of-Attack Total Air Temperature
Left side probes shown. Right side similar.
CS1_CS3_3030_005
Air Data System Probes
Figure 24: Probe Heat System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-57
CS130_30_L2_PHS_ComponentLocation.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
CONTROLS AND INDICATIONS A PROBE HEAT pushbutton annunciator (PBA) powers the ADSP and TAT probe heaters for 2 minutes on the ground provided the auxiliary power unit (APU) is running or external power is connected to the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-58
CS130_30_L2_PHS_ControlsIndications.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
OVERHEAD PANEL
AURAL WARN
PROBE HEAT
INHB
GND ON
L SIDE
OFF
OFF
R SIDE OFF
CS1_CS3_3030_007
OFF
WINDOW HEAT L WSHLD R
Figure 25: Probe Heat Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-59
CS130_30_L2_PHS_ControlsIndications.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
DETAILED DESCRIPTION AIR DATA SYSTEM PROBE Each angle-of-attack (AOA) vane has a vane heater and a case heater. For the left AOA vane, both heaters are powered by DC BUS 1. For the right AOA vane, both heaters are powered by DC BUS 2. The control and distribution cabinets (CDCs) provide power to the heaters. The heaters turn on when electrical power is available on the associated bus, and at least one engine is running or the aircraft is in the air. The air data system probes (ADSPs) use AC power for primary and secondary power sources. Primary AC power is from AC BUS 1, AC BUS 2, or AC ESS BUS. Secondary power for ADSP 1 is AC BUS 2. For ADSP 2, secondary power is from AC BUS 1. The source of secondary power for ADSP 3 and 4 is from static inverters which use 28 VDC power from the DC ESS BUS 3 to provide AC power. AC power from AC BUS 1 and AC BUS 2 is provided through solid-state power controllers (SSPCs). AC power from the AC ESS BUS is through thermal circuit breakers. The static inverters receive 28 VDC through thermal circuit breakers on electrical power control 3 (EPC 3).
When the PROBE HEAT pushbutton annunciator (PBA) is pressed, the white GND ON light illuminates and powers the probe heaters ON for 2 minutes and automatically turns off. The ADS PROBE HEAT GND ON info message is displayed on EICAS when the PROBE HEAT function is operating. The PROBE HEAT function is inhibited on ground unless at least one engine is running and AC power is feeding the aircraft. An ADSP PROBE HEAT FAIL caution message is displayed on EICAS if the associated probe heater is detected failed. The ADSP ISI PROBE HEAT caution message is displayed if both ADSP 3 and ADSP 4 heaters fail. The ADSPs provide data on ADSP and TAT heater operation to EICAS and OMS through the data concentrator unit module cabinets (DMCs).
WARNING
Power for the total air temperature (TAT) heaters is provided through the ADSP, from the AC ESS BUS.
ENSURE PROBE HEAT IS DEACTIVATED BEFORE JACKING THE AIRCRAFT.
The ADSP heaters and TAT probe heaters are switched off whenever the aircraft is on the ground and both engines are off. Otherwise these heaters operate in one of two modes: self-regulating or full mode.
PUTTING THE AIRCRAFT INTO THE AIR MODE AUTOMATICALLY ACTIVATES THE ADSP, TAT AND AOA PROBE HEATERS. ENSURE COVERS ARE REMOVED. HOT PROBES CAN CAUSE INJURY.
The self-regulating mode is enabled once the engines are running, the aircraft is on the ground, and airspeed is less than 55 kt. The probe heat is regulated to maintain a lower probe temperature to avoid overheating. The full mode is enabled when the airspeed is greater than 55 kt or the aircraft is in the air. The ADSP and TAT probe heaters operate in full mode if the controller integral heater fails.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-60
CS130_30_L3_PHS_DetailedDescription.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
CDC 3-6-4 L AOA Vane Heat DC BUS 1 SSPC 15A
LEFT AOA VANE Vane Heater
Logic: Combinational CDC 3-5-2 L AOA Case Heat DC BUS 1 SSPC 10A
Case Heater PROBE HEAT
Logic: Combinational
EICAS
GND ON
CB-G2 10
ADSP 3 ADSP 3A Heat
Heater Power 1
CB-G1 Left TAT Heat
5
LEGEND ADSP AOA CDC DMC TAT
Heater Power
Power
STATIC INVERTER 1
CB-A2 50
LEFT TAT PROBE
Left TAT
LCBP
DC ESS BUS 3
OMS
Inverter 1
28 VDC
115 VAC
Heater Power 2
EPC 3
NOTE • ADSP 3 shown, ADSP 1, 2, and 4 similar. • ADSP 1 primary heater power source is AC BUS 1, and the secondary heater power source is AC BUS 2. • ADSP 2 primary heater power source is AC BUS 2 and the secondary heater power source is AC BUS 1. • ADSP 4 primary heater power source is the AC ESS BUS and the secondary heater power source is DC ESS BUS 3 through static inverter 2. • AOA 2 heater power source is DC BUS 2. • Right TAT probe heater power source is ADSP 4.
Air Data System Probe Angle-of-Attack Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Total Air Temperature ARINC 429 Discrete
CS1_CS3_3031_001
AC ESS BUS
DMCs
Figure 26: Probe Heat Electrical Interface Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-61
CS130_30_L3_PHS_DetailedDescription.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the probe heat system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-62
CS130_30_L2_PHS_MonitoringTests.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
CAS MESSAGES Table 13: ADVISORY Messages
Table 12: CAUTION Messages MESSAGE
LOGIC
MESSAGE
ADS 1 PROBE HEAT FAIL
ADS 1 probe not heated per expectations.
ADS 2 PROBE HEAT FAIL
ADS 2 probe not heated per expectations.
ADS 3 PROBE HEAT FAIL
ADS 3 probe not heated per expectations.
ADS 4 PROBE HEAT FAIL
ADS 4 probe not heated per expectations.
ADS ISI FAIL ADS SAME SOURCE
LOGIC
ADS FAULT
Loss of redundant or non-critical function for the air data system.
ADS 1 DEGRADED
SSEC correction lost and based on default input value(s) for ADS 1 - includes loss of AOA offset.
Both ADS channels of ISI failed (channel 3 and 4).
ADS 2 DEGRADED
Combination of two displays using same source of air data (cover loss of channels 1-2, 1-3, 2-3, 1-4, 2-4).
SSEC correction lost and based on default input value(s) for ADS 2- includes loss of AOA offset.
ADS 3 DEGRADED
SSEC correction lost and based on default input value(s) for ADS 3- includes loss of AOA offset.
ADS 4 DEGRADED
SSEC correction lost and based on default input value(s) for ADS 4- includes loss of AOA offset.
ADS 1 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 2 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 3 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 1 FAIL
Loss of ADS 1 channel.
ADS 2 FAIL
Loss of ADS 2 channel.
ADS 4 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 4 FAIL
Loss of ADS 3 channel.
ADS ISI SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS 3 and 4.
ADS 3 FAIL
Loss of ADS 3 channel.
DUAL ADS FAIL
2 ADS sources are failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Table 14: STATUS Message MESSAGE ADS PROBE HEAT GND ON
TECHNICAL TRAINING MANUAL
For Training Purposes Only
LOGIC PROBE HEAT PBA selected and active.
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30 - Ice and Rain Protection 34-10 Probe Heat System
Page Intentionally Left Blank
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30 - Ice and Rain Protection 34-10 Probe Heat System
Table 15: INFO Messages Table 15: INFO Messages MESSAGE
MESSAGE
LOGIC
34 ADS FAULT - ADS SENSE LINE HEATER 1 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS SENSE LINE HEATER 2 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS SENSE LINE HEATER 3 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS SENSE LINE HEATER 4 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 1 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 2 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 3 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 4 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS 1 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 2 heater or failure/loss of ASDSP 2 OSP electronics.
34 ADS FAULT - ADS 2 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 1 heater or failure/loss of ASDSP 1 OSP electronics.
34 ADS FAULT - ADS 3 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 4 heater or failure/loss of ASDSP 4 OSP electronics.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
34 ADS FAULT - ADS 4 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 3 heater or failure/loss of ASDSP 3 OSP electronics.
34 ADS FAULT - L TAT HEATER INOP
Fuselage TAT not able to be heated on left side. ADS 1 and 3 revert to onside engine TAT (loss of redundancy only).
34 ADS FAULT -R TAT HEATER INOP
Fuselage TAT not able to be heated on right side. ADS 2 and 4 revert to onside engine TAT (loss of redundancy only).
34 ADS FAULT - ADS 1 TAT ELEMENT INOP
ADS 1 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - ADS 2 TAT ELEMENT INOP
ADS 2 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - ADS 3 TAT ELEMENT INOP
ADS 3 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - ADS 4 TAT ELEMENT INOP
ADS 4 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - L AOA VANE INOP
ADS 1 or 3 is reporting a loss or internal failure (HW, SW, strapping changed since power-up) of the LEFT AOA VANE by setting SSM to failure warning or NCD.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_30_L2_PHS_MonitoringTests.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
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30 - Ice and Rain Protection 34-10 Probe Heat System
Table 15: INFO Messages MESSAGE
LOGIC
34 ADS FAULT - R AOA VANE ADS 2 or 4 is reporting a loss or internal INOP failure (HW, SW, strapping changed since power-up) of the LEFT AOA VANE by setting SSM to failure warning or NCD. 34 ADS FAULT - L AOA VANE HEATER INOP
ADS 1 or 3 is reporting a heating failure of the LEFT AOA VANE. This INFO message is inhibited in case the AOA vane heater should not be heated per expectation.
34 ADS FAULT - R AOA VANE ADS 2 or 4 is reporting a heating failure of HEATER INOP the LEFT AOA VANE. This INFO message is inhibited in case the AOA vane heater should not be heated per expectation. 34 ADS FAULT - L AOA VANE DEGRADED
L AOA vane local AOA miscompares with ADSP 1 and ADSP 3 local AOA.
34 ADS FAULT - R AOA VANE L AOA vane local AOA miscompares with DEGRADED ADSP 2 and ADSP 4 local AOA. 34 ADS FAULT - L AOA CASE HEATER INOP
ADS 1 or 3 is reporting that the left AOA case heater current is lower than minimum for operational condition in flight. This INFO message is inhibited in case the AOA case heater should not be heated per expectation.
34 ADS FAULT - R AOA CASE ADS 2 or 4 is reporting that the left AOA HEATER INOP case heater current is lower than minimum for operational condition in flight. This INFO message is inhibited in case the AOA case heater should not be heated per expectation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_30_L2_PHS_MonitoringTests.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
30-41 WINDSHIELD AND SIDE WINDOW HEATING SYSTEM GENERAL DESCRIPTION The windshield and side window heating system (WWHS) applies electrical power to transparent films within the windshields and side windows. The windshields are heated to a higher temperature than the side windows to provide anti-icing and defogging. The side windows are heated to provide defogging only.
The WIPCs provide data on system operation for the engine indication and crew alerting system (EICAS) messages through the data concentrator unit module cabinets (DMCs).
All the windows are monitored for temperature by the associated windshield ice protection controller (WIPC) by using three temperature sensors embedded within each windshield and side window. During normal operation, one sensor provides temperature control and the other two sensors monitor the performance of the first one. In case of a sensor failure, the WIPC switches to one of the remaining two sensors for control and the last one is used as a monitor. Each window contains a transparent heating film, which is embedded between glass laminations. The heating film is configured into three zones. The WIPC supplies 3-phase 115 VAC to the heating film. Each zone is powered by single phase AC power. Electrical power can be turned OFF to each windshield and side window individually using the associated pushbutton annunciator (PBA) on the WINDOW HEAT panel located on the overhead panel. The left windshield WIPC heater power is supplied from the AC BUS 1 and the control power is supplied from DC BUS 1. The right windshield WIPC heater power is supplied from AC BUS 2 and the control power is supplied from DC BUS 2. The left side window WIPC heater power is from the AC ESS BUS and the control power is supplied from DC ESS BUS 3. The right side window WIPC heater power is supplied from AC BUS 2 and the control power is supplied from DC BUS 2. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-68
CS130_30_L2_WSWHS_General.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
LEFT SIDE WINDOW
LEFT WINDSHIELD
2
RIGHT WINDSHIELD
2
RIGHT SIDE WINDOW
2
2
1
1 1
1 Temperature Signal
LEGEND
DMC WIPC
Phase-A Phase-B Phase-C Data Concentrator Unit Module Cabinet Windshield Ice Protection Controller ARINC 429
WIPC 1
3-Phase AC Power
WIPC 3
L SIDE OFF
3X Temperature Sensors DC ESS BUS 3
2
Heating Films
WINDOW HEAT L WSHLD R OFF
OFF
DMC 1
NOTE 1
WIPC 2
CONTROL
AC ESS BUS HEATER
DC BUS 1
AC BUS 1
CONTROL
HEATER
WIPC 4
R SIDE OFF
DMC 2
EICAS OMS
DC BUS 2
AC BUS 2
DC BUS 2
AC BUS 2
CONTROL
HEATER
CONTROL
HEATER
CS1_CS3_3041_001
3-Phase AC Power
Figure 27: Windshield and Window Heat System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-69
CS130_30_L2_WSWHS_General.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
COMPONENT LOCATION The windshield and side window heating system (WWHS) consists of: •
Left and right heated windshield
•
Left and right heated side window
•
Windshield ice protection controllers (WIPCs)
WINDSHIELD There are two electrically-heated windshields located on each side of the forward section of the flight deck. SIDE WINDOW There are two electrically-heated side windows, one located on each side of the flight deck. WINDSHIELD ICE PROTECTION CONTROLLERS Four windshield and ice protection controllers (WIPCs) are installed in the forward equipment bay. Each window is assigned to a dedicated WIPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
FORWARD EQUIPMENT BAY
Side Window
WINDSHIELD ICE PROTECTION CONTROLLERS (4X)
CS1_CS3_3040_003
Windshield
Figure 28: Windshield, Side Window, and Windshield Ice Protection Controllers Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-71
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30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
CONTROLS AND INDICATIONS Each window has its own pushbutton annunciator (PBA). The PBAs are located on the WINDOW HEAT panel in the flight deck overhead panel. The PBAs command their own windshield ice protection controller through the data concentrator unit module cabinets (DMCs) to turn off the windshield/window heat. The OFF legend illuminates on the PBA to confirm that the corresponding window heat function is disabled. Cycling the PBA to OFF then ON again resets the WIPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
OVERHEAD
L SIDE OFF
WINDOW HEAT L WSHLD R OFF
OFF
R SIDE OFF
CS1_CS3_3041_004
WINDOW HEAT PANEL
Figure 29: Window Heat Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-73
CS130_30_L2_WSWHS_ControlsIndications.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
DETAILED DESCRIPTION The windshield and side window heating system (WWHS) applies electrical power to transparent films within the windows. The WWHS is normally on when the aircraft is powered. Window heating can be individually shut off through selection of the associated PBA located on the overhead panel. The controlled temperature of the windshield is set between 45.5°C (114°F) and 51°C (124°F). The side window temperature is set between 36.7° C (98°F) and 43.9°C (111°F). The left side window receives heater power from AC ESS BUS and control power from DC ESS BUS 3. As a result, under emergency conditions where normal AC power is lost, only the left side window can be heated. The WIPCs receive phase of flight data from the data concentrator unit module cabinet (DMC) to moderate power requirements for regulating window heat. For example, at initial power-up in cold soaked conditions, the WIPC regulates window temperature to a lower value to avoid damage to frozen windshields and windows due to thermal stresses. In the event of window heat failure, the WIPC reports through the DMCs, the appropriate WINDSHIELD HEAT FAIL or WINDOW HEAT FAIL caution message for display on EICAS. If the signal from the WINDOW HEAT PBAs is lost or ARINC 429 communications is lost with the DMCs, the WIPC turns on and a WHSLD FAIL ON, or a SIDE WDW HT FAIL ON advisory message is displayed for the affected window. The WIPCs periodically performs built-in tests (BITs) that are capable of detecting electrical power faults, temperature sensor faults, heating film faults, and internal WIPC faults. Test result data is communicated to the onboard maintenance system (OMS) through the DMCs.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-74
CS130_30_L3_WSWHS_DetailedDescription.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
LEFT SIDE WINDOW
LEGEND
DMC WIPC
Phase-A Phase-B Phase-C Data Concentrator Unit Module Cabinet Windshield Ice Protection Controller ARINC 429
2
2
RIGHT WINDSHIELD
2
RIGHT SIDE WINDOW
2
2
1
1
NOTE 1
LEFT WINDSHIELD
1
1
3X Temperature Sensors
Temperature Signal
3-Phase AC Power
Heating Films CB-F1 DC ESS BUS 3
3
L WINDOW CTRL
WIPC 3
WIPC 1
WIPC 2
L CBP
WIPC 4 CDC 4-5-11 R Window CTRL DC SSPC 3A BUS 2
AC ESS BUS
CDC 4-10-1, 2, 3 R Window HTR L SIDE OFF
CDC 3-4-17 L WSHLD CTRL DC SSPC 3A BUS 1 Logic: Always On EPC 1 AC BUS 1
25
L WSHLD HTR
WINDOW HEAT L WSHLD R OFF
OFF
DMC 1
R SIDE OFF
AC BUS 2
SSPC 7.5A
Logic: Always On
DMC 2
CDC 4-4-18 R WSHLD CTRL DC SSPC 3A BUS 2 Logic: Always On
EICAS OMS
R EPC 2 WSHLD HTR 25
AC BUS 2
CS1_CS3_3041_008
CB-H24 L WINDOW HTR 7.5
3-Phase AC Power
Logic: Always On
Figure 30: Windshield and Window Heat System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-75
CS130_30_L3_WSWHS_DetailedDescription.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the windshield and side window heating system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-76
CS130_30_L2_WSWHS_MonitoringTests.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
CAS MESSAGES Table 16: CAUTION Messages MESSAGE
LOGIC
L SIDE WDW HEAT FAIL
Heater for left side window failed.
R SIDE WDW HEAT FAIL
Heater for right side window failed.
L WSHLD HEAT FAIL
Heater for left windshield window failed.
R WSHLD HEAT FAIL
Heater for right windshield window failed.
Table 17: ADVISORY Messages MESSAGE
LOGIC
L WSHLD HEAT FAIL ON
Heater for left windshield failed operative.
R WSHLD HEAT FAIL ON
Heater for right windshield failed operative.
L SIDE WDW HT FAIL ON
Heater for left side window failed operative.
R SIDE WDW HT FAIL ON
Heater for right side window failed operative.
Table 18: STATUS Messages MESSAGE
LOGIC
L SIDE WDW HEAT OFF
Heater for left side window selected OFF.
R SIDE WDW HEAT OFF
Heater for right side window selected OFF.
L WSHLD HEAT OFF
Heater for left windshield window selected OFF.
R WSHLD HEAT OFF
Heater for right windshield window selected OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-77
CS130_30_L2_WSWHS_MonitoringTests.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
30-42 WINDSHIELD WIPERS SYSTEM GENERAL DESCRIPTION Each windshield is provided with a windshield wiper with separate power sources, wiper motors, and associated switches, arms, and wiper blades. The left wiper motor is powered by AC BUS 1 and the right wiper motor is powered by AC BUS 2.
The wiper motors run continuous built-in tests (CBIT). In the event of failure the electronic controller sends failure information through the data concentrator unit module cabinets (DMCs) to the onboard maintenance system (OMS). There are no crew alerting system (CAS) messages for the windshield wiper system.
The wiper motors have electronic controllers and are electrically connected to each other for synchronization. This allows them to move together when both windshield wiper switches are selected to the same speed. The synchronization feature does not prevent the other wiper from operating if one wiper motor fails. The wiper controller incorporates a thermal protection that removes power to the motor if the motor overheats. After a cool down period, the controller reapplies power to the motor. When not in operation, the wipers move to the parked position, resting vertically at the center post of the windshield frame. If a wiper motor fails, it moves to the park position. The windshield wiper operation can be reset by setting the rotary switch to OFF and back to any of the other positions. The winshield wiper switch has the following selections: •
OFF: The onside wiper moves to the vertical parked position
•
INT: One wiping cycle every 5 seconds
•
SLOW: 80 cycles per minute
•
FAST: 120 cycles per minute
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-78
CS130_30_L2_WWS_General.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
Right Wiper Arm
Park Position
RIGHT WIPER MOTOR ELECTRONIC CONTROLLER CDC 4-9-4 R Wiper Motor AC SSPC 5A BUS 2 Logic: Always On
Left Wiper Arm
LEFT WIPER MOTOR sync
ELECTRONIC CONTROLLER
DMC 1 and DMC 2
OFF
LEGEND DMC OMS
Data Concentrator Unit Module Cabinet Onboard Maintenance System ARINC 429 Discrete
Logic: Always On
WIPER OFF
INT SLOW
OMS
FAST
INT SLOW FAST
CS1_CS3_3042_002
WIPER
CDC 3-9-6 L Wiper Motor AC SSPC 5A BUS 1
Figure 31: Windshield Wiper System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-79
CS130_30_L2_WWS_General.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
COMPONENT LOCATION The windshield wiper system consists of the following: •
Wiper motors
•
Wiper arms
•
Wiper blades
WIPER MOTOR There are two wiper motors. Each motor is located below its windshield. WIPER ARM There are two wiper arms. Each wiper arm is attached to its windshield wiper motor. WIPER BLADE There are two wiper blades. A wiper blade is attached to each wiper arm.
CAUTION To avoid damage to the windshield do not operate windshield wipers on dry glass.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-80
CS130_30_L2_WWS_ComponentLocation.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
Wiper Blade
Wiper Arm
CS1_CS3_3042_003
Wiper Motor
Figure 32: Windshield Wiper Motor, Wiper Arm, and Wiper Blade Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-81
CS130_30_L2_WWS_ComponentLocation.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
CONTROLS AND INDICATIONS The WIPER selector switches are located on the left and right wiper panel on the overhead panel of the flight deck. The switches allow selection of three speeds: intermittent (INT), SLOW, and FAST.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-82
CS130_30_L2_WWS_ControlsIndications.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
OVERHEAD PANEL
READING
OFF
BRT
READING
WIPER OFF
WIPER OFF
INT SLOW FAST
OFF
BRT
INT SLOW
CS1_CS3_3042_004
FAST
Figure 33: Windshield Wipers System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-83
CS130_30_L2_WWS_ControlsIndications.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
Page Intentionally Left Blank
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30-84
CS130_30_ILB.fm
ATA 36 - Pneumatic
BD-500-1A10 BD-500-1A11
36 - Pneumatic
Table of Contents 36-10 Bleed Air System ..........................................................36-2 General Description .........................................................36-2 Left and/or Right Engine Bleed Air...............................36-2 Auxiliary Power Unit Bleed Air .....................................36-2 External Air Source ......................................................36-2 Component Location .......................................................36-4 Precooler Exhaust Door...............................................36-4 Bleed Temperature Sensor..........................................36-4 Bleed Monitoring Pressure Sensor ..............................36-4 Precooler......................................................................36-4 Fan Air Valve ...............................................................36-4 Intermediate Pressure Check Valve ............................36-4 High-Pressure Valve ....................................................36-4 Pressure-Regulating and Shutoff Valves .....................36-4 Cross-Bleed Valve .......................................................36-6 High-Pressure Ground Connection..............................36-6 Auxiliary Power Unit Check Valve................................36-6 APU Bleed Air Valve ....................................................36-8 Integrated Air System Controllers ..............................36-10 Detailed Component Information ....................................36-12 Integrated Air System Controller ...............................36-12 Electrical Interface .....................................................36-14 Controls and Indications .................................................36-16 Controls .....................................................................36-16 Indications .................................................................36-18 Detailed Description ......................................................36-20 Engine Bleed Air 4th and 8th Bleed Port Switching ..............................36-20 Pressure-Regulating Shutoff Valve Operation ...........36-22 Fan Air Valve Operation ............................................36-24 Precooler Exhaust Door.............................................36-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Bleed Air Source Selection ...................................... 36-28 Cross-Bleed Valve Operation.................................... 36-28 Monitoring and Tests .................................................... 36-30 CAS Messages ......................................................... 36-31 Practical Aspects ............................................................ 36-34 High-Pressure Valve, Pressure-Regulating Shutoff Valve, and Fan Air Valve Deactivation ......... 36-34 Cross-Bleed Valve Deactivation................................ 36-36 Precooler Exhaust Door Manual Opening and Closing.................................... 36-38 PCE Door Test .......................................................... 36-40 Onboard Maintenance System Tests ........................ 36-42 36-21 Bleed Air Leak and Overheat Detection System ....... 36-44 General Description ...................................................... 36-44 Bleed Air Leak and Overheat Detection System Operation .................................... 36-46 Component Location ..................................................... 36-48 Integrated Air System Controllers ............................. 36-48 Dual-Detection Loops................................................ 36-50 Manifolds................................................................... 36-50 Detailed Component Information ................................... 36-52 Detection-Loops ........................................................ 36-52 Manifolds................................................................... 36-52 Controls and Indications ................................................ 36-54 Detailed Description ...................................................... 36-56 Bleed Air Leak and Overheat Detection System Activation ..................................... 36-56 BALODS Leak Localization ...................................... 36-58 Monitoring and Tests ..................................................... 36-60 CAS Messages ......................................................... 36-61
TECHNICAL TRAINING MANUAL
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36-i
CS130_M5_36_PneumaticTOC.fm
36 - Pneumatic
List of Figures Figure 1: Bleed Air System Schematic .................................36-3 Figure 2: Engine Bleed Air System Components .................36-5 Figure 3: APU Check Valve, Cross-Bleed Valve, and High-Pressure Ground Connection Location.................................................................36-7 Figure 4: APU Bleed Air Valve .............................................36-9 Figure 5: Integrated Air System Controllers (IASCs) ..........36-11 Figure 6: Integrated Air System Controller ........................36-13 Figure 7: Electrical Interface...............................................36-15 Figure 8: AIR Panel ............................................................36-17 Figure 9: Bleed Air System Indications...............................36-19 Figure 10: Engine Bleed Port Switching ...............................36-21 Figure 11: PRSOV Operation ...............................................36-23 Figure 12: Fan Air Valve Operation ......................................36-25 Figure 13: Precooler Exhaust Door Operation .....................36-27 Figure 14: Bleed Air Source Selection and Cross-Bleed Valve Operation..............................36-29 Figure 15: High-Pressure Valve, Pressure-Regulating Shutoff Valve, and Fan Air Valve Deactivation ...........................36-35 Figure 16: Cross-Bleed Valve Deactivation..........................36-37 Figure 17: PCE Exhaust Door Manual Opening and Closing ..............................36-39 Figure 18: Precooler Exit Door Test .....................................36-41 Figure 19: Bleed Air System Initiated Built-In Test ...............36-43 Figure 20: BALODS Operation .............................................36-45 Figure 21: BALODS Zones...................................................36-47 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 22: Integrated Air System Controller......................... 36-49 Figure 23: Bleed Air Leak and Overheat Detection System Components .......................... 36-51 Figure 24: Manifold .............................................................. 36-53 Figure 25: Bleed Air Leak and Overheat Detection System Indications ............................. 36-55 Figure 26: Bleed Air Leak and Overheat Detection System Operation............................... 36-57 Figure 27: BALODS Leak Localization ............................... 36-59
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-ii
CS130_M5_36_PneumaticLOF.fm
36 - Pneumatic
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-iii
CS130_M5_36_PneumaticLOF.fm
PNEUMATIC - CHAPTER BREAKDOWN CHAPTER BREAKDOWN
Bleed Air System
1 Bleed Air Leak and Overheat Detection System
2
36 - Pneumatic 36-10 Bleed Air System
36-10 BLEED AIR SYSTEM GENERAL DESCRIPTION The aircraft bleed air system (BAS) receives hot bleed air from the engines, auxiliary power unit (APU), or an external air source. Pressurized air from these sources enters common ducting in the mid fuselage belly before being distributed to the user systems. A crossbleed valve (CBV) located in the manifold, isolates left and right side bleed air systems. Opening the CBV interconnects the two sides. The integrated air system controllers (IASCs) control and monitor the BAS. Each IASC is powered by two DC ESS BUSES. LEFT AND/OR RIGHT ENGINE BLEED AIR
valve (FAV) modulates the amount of engine fan air passing through the precooler in order to maintain the bleed air temperature limits. A bleed temperature sensor (BTS) provides the IASCs with temperature information to control the FAV.The fan air exhausts into the engine core compartment where it exhausts overboard. A precooler ex-haust (PCE) door opens automatically under certain high bleed air demand conditions to increase cool-ing airflow. Bleed air, downstream of the bleed air supply of both engines, is directed to a common high-pressure duct. A cross-bleed valve (CBV), mounted on the high-pressure duct, connects and isolates bleed air from both engines. The PRSOV and FAV are commanded closed by the IASC when the associated ENG FIRE PBA is pressed. Each valve has a thermal fuse that cuts its power if the temperature exceeds 250°C.
The left and right engines are the normal source of bleed air in flight. The BAS uses intermediate pressure bleed air from the 4th stage engine compressor, or high-pressure 8th stage bleed air supplied through the high-pressure valve (HPV). An intermediate pressure check valve (IPCKV) prevents 8th stage air from back flowing into the 4th stage port when the HPV is open.
AUXILIARY POWER UNIT BLEED AIR
Switching between 4th and 8th stage engine bleed air is based aircraft configuration and demand as calculated by the electronic engine control (EEC). The IASC uses this information to control the operation of the high-pressure valve (HPV). A bleed monitoring pressure sensor (BMPS) sends pressure information to the IASC for HPV position monitoring.
When the engines are started, the APU remains the primary bleed source to optimize engine performance during takeoff. When the APU performance decreases with altitude, the IASC automatically closes the APU BAV and switches the bleed source to the engines. When the APU bleed is selected off, the APU check valve prevents engine bleed air from reaching the APU.
Downstream of the BMPS, a pressure-regulating and shutoff valve (PRSOV) regulates bleed air pressure according to the bleed air demand requirements to a maximum of 59 psi. The torquemotor controlled PRSOV has a full closed microswitch for valve position reporting to the IASCs. The air conditioning pack inlet pressure sensors (PIPSs), (ATA 21), are used by the IASCs to regulate the pressure downstream of the PRSOV. A precooler, located downstream of the PRSOV control the engine bleed air temperature using engine fan air. A torquemotor controlled fan air Copyright © Bombardier Inc. CS130-21.07-06-00-121418
During normal ground operation, the IASC defaults the bleed air source to the APU. The APU bleed air valve (BAV) is controlled by the APU electronic control unit (ECU), (ATA 49). The IASC provides a load demand signal to the APU ECU. The ECU adjusts the APU load compressor output to meet the BAS demand.
EXTERNAL AIR SOURCE On the ground, the BAS can be supplied with pressurized air from an external air source, through the high-pressure ground connection (HPGC). The HPGC check valve closes when not in use to prevent bleed air system loss.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-2
CS130_36_L2_BAS_General.fm
36 - Pneumatic 36-10 Bleed Air System
L ENG
L EEC DC ESS 1
FIRE
DC ESS 2
IASC 1
DMC 1 and DMC 2
L PIPS 1 L PIPS 2
ENGINE AND APU FIRE PANEL
EICAS SYN OMS
T/M
Fan
FAV BMPS
T/M
PS
Precooler
To Right WAIS To Right Pack
To FTIS To Left Pack
BTS
M
From Right Engine Bleed Air System
4th Stage IPCKV 8th Stage
PRSOV
CROSS-BLEED VALVE
SOL
HPV
APU ECU
Overboard
APU CHECK VALVE
To Air Turbine Starter
BMPS BTS DMC ECU EEC FAV FTIS HPGC HPV
Fan Air Bleed Air Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Data Concentrator Unit Module Cabinet Electronic Control Unit Electronic Engine Control Fan Air Valve Fuel Tank Inerting System High Pressure Ground Connection High-Pressure Valve
IASC IPCKV PIPS PRSOV WAIS
Integrated Air System Controller Intermediate Pressure Check Value Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System ARINC 429
APU BLEED AIR VALVE
SOL
External Air
LEGEND
HPGC
APU
Motor
NOTE
SOL
Solenoid
Left Bleed Air System shown. Right Bleed Air System similar.
T/M
Torque Motor
M
CS1_CS3_3600_013
To Left WAIS
LEFT ENGINE
Figure 1: Bleed Air System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-3
CS130_36_L2_BAS_General.fm
36 - Pneumatic 36-10 Bleed Air System
COMPONENT LOCATION
BLEED MONITORING PRESSURE SENSOR
The main components of the bleed air system include:
The bleed monitoring pressure sensor (BMPS) is located in the pylon area. It is connected to the HP air duct downstream of the high-pressure shutoff valve, by a sense line.
•
Precooler exhaust (PCEs) doors
•
Bleed temperature sensors (BTSs)
•
Bleed monitoring pressure sensors (BMPSs)
•
Precoolers
•
Fan air valve (FAV)
•
Intermediate pressure check valve (IPCKV)
•
High-pressure valves (HPVs)
•
Pressure-regulating shutoff valves (PRSOVs)
•
Cross-bleed valve (CBV) (Refer to figure 3)
INTERMEDIATE PRESSURE CHECK VALVE
•
High-pressure ground connection (HPGC) (Refer to figure 3)
•
APU check valve (Refer to figure 3)
An intermediate pressure check valve (IPCKV) is located on the engine bleed air duct, at the 1 o’clock position.
•
APU bleed air valve (BAV) (Refer to figure 4)
HIGH-PRESSURE VALVE
•
Integrated air system controllers (IASCs) (Refer to figure 5)
The high-pressure valve (HPV) is located on the HP air duct, at the 2 o’clock position. It is also referred to as the high-pressure shutoff valve (HPSOV).
PRECOOLER The precooler, also known as the precooler exchanger is located on the engine pylon. The engine needs to be lowered in order to replace the precooler. FAN AIR VALVE The fan air valve (FAV) is located upstream of the precooler.
PRECOOLER EXHAUST DOOR The precooler exhaust (PCE) door is located on the engine right core cowl inner fixed structure, at the 3 o’clock position, aft of the oil tank access door.
PRESSURE-REGULATING AND SHUTOFF VALVES The pressure-regulating and shutoff valve (PRSOV) is located on the engine bleed air duct downstream of the HPV at the 1 o’clock position.
BLEED TEMPERATURE SENSOR The bleed temperature sensor is located on the high-pressure air duct, downstream of the precooler in the pylon area.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-4
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
Bleed Temperature Sensor
Bleed Monitoring Pressure Sensor (mounted on pylon)
Precooler Pylon
Fan Air Valve
Intermediate Pressure Check Valve
ENGINE RIGHT CORE COWL
High-Pressure Pressure-Regulating Valve Shutoff Valve
RIGHT ENGINE (OUTBOARD VIEW)
CS1_CS3_3615_002
Precooler Exhaust Door (access panel removed for clarity)
Figure 2: Engine Bleed Air System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-5
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
CROSS-BLEED VALVE The cross-bleed valve (CBV) is located in the mid forward section of the wing-to-body fairing (WTBF) area. HIGH-PRESSURE GROUND CONNECTION The high-pressure ground connection (HPGC) is located in the left forward section off the WTBF area. AUXILIARY POWER UNIT CHECK VALVE The auxiliary power unit (APU) check valve is located in the right aft section of the WTBF area.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-6
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
A B C
CROSS-BLEED VALVE
B HIGH-PRESSURE GROUND CONNECTION
C APU CHECK VALVE
CS1_CS3_3615_010
A
Figure 3: APU Check Valve, Cross-Bleed Valve, and High-Pressure Ground Connection Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-7
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
APU BLEED AIR VALVE The APU bleed air valve (BAV) is located on the APU.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-8
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
CS1_CS3_3615_015
AUXILIARY POWER UNIT (BOTTOM VIEW)
APU BLEED AIR VALVE
Figure 4: APU Bleed Air Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-9
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
INTEGRATED AIR SYSTEM CONTROLLERS Two integrated air system controllers (IASCs), IASC 1 and IASC 2 are located in the mid equipment bay forward rack.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-10
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
IASC 1 LEGEND IASC
Integrated Air System Controller
MID EQUIPMENT BAY FWD RACK
CS1_CS3_3615_017
IASC 2
Figure 5: Integrated Air System Controllers (IASCs) Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-11
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
DETAILED COMPONENT INFORMATION INTEGRATED AIR SYSTEM CONTROLLER Each IASC incorporates a channel A and a channel B. Channel A is the primary control channel, and channel B is the standby channel. On the ground the IASC 1 alternates with IASC 2 based on odd and even days. If a failure is detected on the primary control channel, the IASC switches to the healthy channel. For safety purposes, a third independent and dissimilar channel is integrated and referred to as the safety channel C. This allows system control independence from system failures annunciation. IASC channel functions are assigned accordingly: •
The left and right bleed pressure and temperature control is performed by the associated IASC primary control channel A or channel B
•
The left and right bleed pressure monitoring is performed by the associated IASC primary control channel A and channel B
•
The left and right bleed temperature monitoring is performed by the associated IASC channel C, and the opposite IASC channel C. Temperature monitoring is also performed by all channels for auto-shutdown purposes
•
The CBV control and monitoring is performed by the left IASC channel B, and the right IASC channel B
•
The APU bleed control is performed by the left and right IASC primary control channel A or channel B
An initiated built-in test (IBIT) of the bleed air system (BAS) can be performed from the onboard maintenance system (OMS) for each channel and controller.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-12
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
IASC 1
Channel A
Channel B
Safety Channel C
• Left engine bleed management control and monitoring
• Left engine bleed management control and monitoring
• Left and right engine bleed management monitoring
• Left engine bleed shutoff control and monitoring
• Left engine bleed shutoff control and monitoring
• Left and right temperature monitoring
• APU bleed control
• APU bleed control • Cross-bleed control and monitoring
EICAS Data Exchange
DMC
SYN OMS
IASC 2
Channel A
Channel B
Safety Channel C
• Right engine bleed management control and monitoring
• Right engine bleed management control and monitoring
• Left and right engine bleed management monitoring
• Right engine bleed shutoff control and monitoring
• Right engine bleed shutoff control and monitoring
• Left and right temperature monitoring
• APU bleed control
• APU bleed control CS1_CS3_3620_004
• Cross-bleed control and monitoring
LEGEND ARINC 429
Figure 6: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-13
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
ELECTRICAL INTERFACE Each IASC is electrically powered by two independent DC ESS BUSES: •
IASC 1 is powered by DC ESS BUS 1 and DC ESS BUS 2
•
IASC 2 is powered by DC ESS BUS 2 and DC ESS BUS 3
The IASCs power the high-pressure valve (HPV), fan air valve (FAV), cross-bleed valve (CBV), and the pressure-regulating shutoff valve (PRSOV). The control and distribution cabinet (CDC) provides the signal to enable PRSOV operation through the IASC. The CDC command is provided when the AIR panel BLEED PBA is selected and a minimum of 15 psi bleed air is available. The PRSOV is monitored by the pack inlet pressure sensors (PIPS). When the PRSOV is open, the PIPS provide pressure information to control the output of the PRSOV. The PIPS also provide monitoring for low-pressure and overpressure conditions. The PRSOV is monitored for the closed position by a microswitch. The HPV solenoid operated valve is controlled by the IASC. The opening and closing of the HPV is based on intermediate pressure temperature information calculcated by the electronic engine control (EEC) and the bleed monitoring pressure sensors (BMPSs). The BMPS, located downstream of the intermediate pressure port check valve is used to monitor the position of the HPV.
When a main engine start signal is received when the aircraft is on the ground: the following valves are closed for the duration of the start plus 30 seconds. •
Wing anti-ice valve (WAIV)
•
Left and right flow control valves (FCV)
•
Trim air pressure regulating valve (TAPRV)
•
Trim air shutoff valve (TASOV)
When the engine start is complete, the valves reopen. In flight, the 30 second delay is not used and the valves open immediately after the engine starts. The EEC provides the following information: •
Engine running
•
Main engine start
•
N2
•
Intermediate pressure port temperature
The EEC information is used to configure the bleed air system and control system operation.
The PIPS and BMPS are powered from the DC ESS BUSES. The bleed temperature sensors (BTS) provide temperature information to the IASC for monitoring and control of the FAV. The IASC provides the required pneumatic demand information to the APU electronic control unit (ECU), when the APU is supplying bleed air. The APU load compressor provides the required output through the APU bleed air valve (BAV). The APU BAV is powered by the APU ECU.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-14
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
EPC 2 DC ESS BUS 2
IASC 2 IASC 2A
15
CH 2A
R CBP DC ESS BUS 3 CDC 1-6-10 L ECS SNSR DC ESS SSPC 3A BUS 1
15
L BTS IASC 2B
CH 2B CH S
L BLEED TEMPERATURE SENSOR
IASC 1 L EEC L PACK INLET PRESSURE SENSOR 1
L PIPS 2
L PACK INLET PRESSURE SENSOR 2
L PIPS 2 APU Bleed Enable
L BLEED MONITORING PRESSURE SENSOR
L BMPS
L BTS
Logic: Always On
Main Engine Start N2 Engine Running Intermediate Pressure Port Temperature
APU ELECTRONIC CONTROL UNIT
APU BLEED AIR VALVE
EPC 2 3
L BLEED SNSR
Main Engine Start Signal
Close Signal To: • WAIV’s • FCV’s • TAPRV • TASOV LEGEND
CDC 1 L FAV Torquemotor L PRSOV CMD
L FAN AIR VALVE
L PRSOV Enable L PRSOV Torquemotor L PRSOV Closed
L PRSOV
EPC 1 CBP DC ESS BUS 1
15
IASC 1A
EPC 2 CBP NOTE
DC ESS BUS 2
15
IASC 1B
CH 1A
L HPV SOL CMD
CBV Open/Closed CMD CBV Open CH 1B CBV Closed CH S
HIGH-PRESSURE VALVE
CROSS-BLEED VALVE
Left Bleed Air System Shown. Right Bleed Air System Similar.
BMPS BTS CBV CDC CL/OP CMD EEC FAV FCV FC FO HPV IASC PRSOV SNSR SOL TAPRV TASOV WAIV
Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Close/Open Command Electronic Engine Control Fan Air Valve Flow Control Valve Full Close Full Open High-Pressure Valve Integrated Air System Controller Pressure-Regulating Shutoff Valve Sensor Solenoid Trim Air Pressure Regulating Valve Trim Air Shutoff Valve Wing-Anti-Ice Valve ARINC 429
CS1_CS3_3611_005
DC ESS BUS 2
Figure 7: Electrical Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-15
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
CONTROLS AND INDICATIONS CONTROLS The bleed air system (BAS) is controlled from the AIR panel, located on the flight deck overhead panel. Bleed air valve (BAV) operation is controlled by pushbutton annunciators (PBAs) and a rotary selector. When the L BLEED PBA, the R BLEED PBA, and the APU BLEED PBA are in the ON position, the operation of the associated BAV is automatically managed by the IASCs. When a PBA is selected OFF, the valve is forced closed and the OFF legend illuminates.
The cross-bleed valve (CBV) is controlled by the rotary XBLEED switch. The CBV is controlled automatically by the integrated air system controller (IASC) when the XBLEED switch is in the AUTO position. The IASC opens the CBV when there is a single source of bleed air supply (one engine, APU or external air). The CBV is closed or open when the XBLEED switch is selected to MAN CLSD or MAN OPEN respectively. An ENG BLEED MISCONFIG caution message is displayed if the CBV is open with both engine bleeds on.
The amber L (R) BLEED PBA FAIL light illuminates for the following conditions: •
Overpressure
•
Low pressure
•
Loss of both BTS sensing elements
•
HPV failed open
•
PRSOV failed open
•
IASC 1 or IASC 2 failure
The corresponding L (R) BLEED FAIL caution message is displayed on the engine indication and crew alerting system (EICAS) page. An amber FAIL light illuminates in the APU BLEED PBA in the event of failure to supply APU bleed air or the loss of bleed leak detection capability. The APU BLEED FAIL caution message is displayed on EICAS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-16
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
RAM AIR OVERHEAD PANEL
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
OPEN
FAIL OFF R BLEED
APU BLEED
FAIL OFF
FAIL OFF
CS1_CS3_3620_002
BLEED AIR PANEL
Figure 8: AIR Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-17
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
INDICATIONS The bleed air system (BAS) operation indication is displayed on the AIR synoptic page. Bleed air flow lines, actual duct temperature, valve position in normal operation, valve position in fail operation, and pressure display with symbols and color codes. The EXT AIR symbol is displayed on the AIR synoptic page when the air pressure is detected, and the engines and APU bleed air are not operating.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-18
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
FLOW LINES
STATUS
AIR IR
DOOR OR
ELECC
FLT CTR CTRLL
FUEL FUE EL
HYD YD
AVIONIC VIONIC
INFO
CB
Condition Normal flow. No flow. Abnormal flow (overheat, out of range pressure,leak).
VALVE POSITION NORMAL OPERATION Symbol
Condition
Symbol
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°CC
10 °C °
15 °CC
Open with flow.
Open with flow.
Open with no flow.
Open with no flow.
5 PSI
40 PSI
-PSI
Normal.
EXT AIR
R PACK
40
XBLEED
Left Pressure-Regulating Shutoff Valve
EXT AIR
Condition Displayed when pressure detected and engine/APU bleeds are off.
HI Right Bleed Air Pressure
PSI
DECLUTTERED SYMBOLOGY LOGIC
Abnormal.
HI
Left Bleed Air Pressure
Invalid.
Symbol
RAM AIR
Condition Closed.
Condition
15 °CC
TRIM AIR L PACK
Symbol
AFT
CARGO
VALVE POSITION FAIL OPERATION
Closed.
PRESSURE
FWD
CKPT 23 °C 22 °C
APU
Invalid. SYNOPTIC PAGE – AIR
40 PSI
Right Pressure-Regulating and Shutoff Valve
CS1_CS3_3620_005
Symbol
Figure 9: Bleed Air System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-19
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
DETAILED DESCRIPTION The active integrated air system controller (IASC) automatically performs engine bleed air 4th and 8th bleed port switching, and controls the operation of the pressure-regulating shutoff valve (PRSOV), when the associated BLEED PBA on the AIR panel is selected ON.
To switch back to the engine IP port, the IASC uses the IPP information from the EEC. Once the actual pressure at the IP port matches the IPP calculated pressure, the IASC closes the HPV. During, descent, and approach or steep approach, the HP/IP threshold is increased to prevent excessive bleed port switching.
ENGINE BLEED AIR 4TH AND 8TH BLEED PORT SWITCHING
Temperature
The high-pressure (HP)/intermediate-pressure (IP) switching selects the bleed air from the engine IP or HP port based on the system configuration (number of packs, WAIS on) and demand of the user systems. The IASC automatically selects the proper engine bleed port based on pressure requirements and temperature limitations.
The bleed port switching logic also considers the IP temperature (IPT) to maintain WAIS performance under single or dual bleed source operation. If the IPT is lower than the WAIS threshold of 230°C (446°F), the IASC commands the HPV open. The HPV closes if the IPT increases above 240°C (464°F),
Pressure requirements are based on the required bleed flow, user system configuration, and engine thrust settings as calculated by the electronic engine control (EEC).
To protect the bleed air system downstream of the engine ports, the IASC inhibits the IP/HP switching when the HP port temperature (HPT) is above the maximum temperature threshold of 490°C (914°F). The HPV closes if the HPT increases above
Temperature limitations are based on the maximum allowable system temperature, and the minimum temperature to provide satisfactory WAIS operation.
If the HPV is open, and the maximum temperature threshold of 500°C (932°F) is reached, the IASC closes the HPV.
The engine bleed port switching is done by electrical control of the HPV. Pressure The IASC opens the HPV based on the EEC intermediate pressure port (IPP) calculated pressure information. The EEC calculation is the minimum pressure necessary to provide the required total bleed airflow in the system. If the actual pressure at the IP port, as indicated by the BMPS goes below a specified threshold, the IASC opens the HPV and the HP port supplies the pressure. The threshold varies based on the system configuration and demand. The HPV downstream pressure is limited to 75 +/- 5 psi.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-20
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L ENG
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
FIRE
OPEN
R PACK XBLEED
LEGEND
L EEC
DMC 1 and DMC 2
FAIL OFF
ENGINE AND APU FIRE PANEL
EICAS
R BLEED APU BLEED
ACS APU ATS BMPS BTS CBV CDC CMD DMC FAV FTIS HPGC HPV IASC PIPS PRSOV WAIS
SYN
FAIL OFF
FAIL
OMS
OFF
AIR PANEL CDC 1
IASCs • WAIS • ACS
L PIPS 1 L PIPS 2
L PRSOV CMD
LEFT ENGINE T/M
Fan
M
Users • L ACS • L WAIS • FTIS
FAV
BMPS
PS
T/M
Precooler
BTS
SOL
Motor
SOL
Solenoid
T/M
Torque Motor
M From Right Engine Bleed Air System
4th Stage 8th Stage
Fan Air Bleed Air Air Conditioning System Auxiliary Power Unit Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Command Data Concentrator Unit Module Cabinet Fan Air Valve Fuel Tank Inerting System High-Pressure Ground Connection High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System AFDX ARINC 429
PRSOV
CBV
To ATS
APU BLEED AIR VALVE
SOL
NOTE
External Air
Left Bleed Air System shown. Right Bleed Air System similar.
HPGC
APU
CS1_CS3_3600_019
HPV
Figure 10: Engine Bleed Port Switching Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-21
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
PRESSURE-REGULATING SHUTOFF VALVE OPERATION
•
When the BLEED PBA is selected ON, the control and distribution cabinet (CDC) provides an enable signal to the IASC, allowing the PRSOV to open and supply engine bleed air.
The BTS detects an overtemperature (refer to Fan Air Valve Operation)
•
The bleed air leak and overheat detection system (BALODS) detects a bleed air leak in the downstream ducting
Once enabled, PRSOV operation is controlled by the IASC. The IASC uses the pressure signal supplied by the air conditioning pack inlet pressure sensors (PIPSs), to maximize air conditioning performance. The IASC modulates the torque motor of the PRSOV to provide bleed air at a pressure of 41.5 psi. Refer to ATA 21, Environmental Control for additional information on PIPS.
If the PRSOV fails in the open position, the amber FAIL light illuminates in the associated L (R) PBA.
During a cross-bleed main engine start, the IASC controls the operating engine PRSOV to maintain 45 psi for engine starting. The PIPSs on the operating engine side provides the pressure signal to the IASC. When the EEC signals that the engine is not running, the IASC closes the associated PRSOV and HPV. The PRSOV is commanded closed by any of the following conditions: •
BLEED PBA is selected OFF
•
During a start of the associated engine. The PRSOV remains closed for the duration of the start plus an additional 30 seconds during a ground start. In flight, the 30 second delay is removed.
•
No bleed air flow demand (no WAI demand or air conditioning demand), provided the RAM AIR PBA is not selected open. If the RAM AIR PBA is selected open, the PRSOV is prevented from closing in order to provide trim air to the trim air system. Refer to ATA 21, Air Conditioning for more information
•
The associated ENG FIRE PBA is selected. This signal is provided through the data concentrator unit module cabinets (DMCs)
•
Overpressure of 45 psi (if bleed demand) or 60 psi (if no bleed demand), for more than 15 seconds, as reported by the associated PIPS. This protects the bleed air ducting from high engine bleed pressure
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-22
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L ENG
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
FIRE
OPEN
R PACK XBLEED
LEGEND
L EEC
DMC 1 and DMC 2
FAIL OFF
ENGINE AND APU FIRE PANEL
EICAS
R BLEED APU BLEED
ACS APU ATS BMPS BTS CBV CDC CMD DMC FAV FTIS HPGC HPV IASC PIPS PRSOV WAIS
SYN
FAIL OFF
FAIL
OMS
OFF
AIR PANEL CDC 1
IASCs • WAIS • ACS
L PIPS 1 L PIPS 2
L PRSOV CMD
LEFT ENGINE T/M
Fan
M
Users • L ACS • L WAIS • FTIS
FAV
BMPS
PS
T/M
Precooler
BTS
SOL
Motor
SOL
Solenoid
T/M
Torque Motor
M From Right Engine Bleed Air System
4th Stage 8th Stage
Fan Air Bleed Air Air Conditioning System Auxiliary Power Unit Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Command Data Concentrator Unit Module Cabinet Fan Air Valve Fuel Tank Inerting System High-Pressure Ground Connection High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System AFDX ARINC 429
PRSOV
CBV
To ATS
APU BLEED AIR VALVE
SOL
NOTE
External Air
Left Bleed Air System shown. Right Bleed Air System similar.
HPGC
APU
CS1_CS3_3600_019
HPV
Figure 11: PRSOV Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-23
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
FAN AIR VALVE OPERATION The IASC uses feedback from the BTS to control the fan air valve (FAV) in order to maintain a BAS temperature range between 170°C (338°F) and 232°C (450°F). The temperature regulation setpoint is determined by the IASC based on the number of engines operating, the number of air conditioning packs operating, and wing anti-ice system (WAIS) operation. The nominal setpoint with air conditioning running is 200°C (392°F). With the L (R) BLEED PBA selected on and the associated engine operating, the IASC enables the FAV to open. When the engine bleed air supply is commanded closed automatically or the L (R) BLEED PBA is selected OFF, and the PRSOV is confirmed closed by the IASC, the FAV closes. Should the PRSOV fail in the open position, the FAV to stay open to allow fan air cooling to the precooler. The FAV closes simultaneously with the PRSOV when the associated ENG FIRE PBA is selected. If the BTS detects a temperature greater than 257°C (495°F) for more than 30 seconds, the L (R) BLEED OVHT caution message is displayed on EICAS and the FAIL light on the L (R) BLEED PBA illuminates. The IASC closes the associated PRSOV (and HPV). The associated BLEED PBA on the AIR panel is selected OFF to match the valve position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-24
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L PACK FAIL OFF
MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
LEGEND
R PACK XBLEED
FAIL OFF R BLEED
APU BLEED
FAIL OFF
L ENG
FAIL OFF
FIRE
AIR PANEL
EICAS DMCs ENGINE AND APU FIRE PANEL
OMS
CDC 1
ATS BMPS BTS CDC CMD DMC FAV HPV IASC PIPS PRSOV WAIS
Fan Air Bleed Air Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Control and Distribution Cabinet Command Data Concentrator Unit Module Cabinet Fan Air Valve High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System AFDX ARINC 429
L PRSOV CMD
IASCs
LEFT ENGINE T/M
FAV BMPS
PS
T/M
Precooler
To Bleed Air System
4th Stage 8th Stage
BTS
SOL
PRSOV
HPV
NOTE To ATS
Left Bleed Air System shown. Right Bleed Air System similar.
CS1_CS3_3600_020
Fan
Figure 12: Fan Air Valve Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-25
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
PRECOOLER EXHAUST DOOR The precooler exhaust (PCE) door is located on the right inner fixed structure (IFS) panel at the 3 o'clock position. The door actuator is controlled by channel A of the electronic engine control (EEC). Operation of the actuator unlatches the exhaust door which is pushed open by two pistons. The PCE door can be opened electrically by the EEC, or manually by rotating the latches in a counterclockwise direction using a 0.635 cm (0.250 in.) drive tool. The door can only be closed manually. The PCE door has square drive adaptors to operate the upper and lower pistons, actuator, and striker latch. When closing the PCE door, an actuator reset check hole is used to confirm the actuator is in the correct position before securing the upper and lower pistons. When the precooler demand is high, the door opens based on logic from the EEC. The door actuator receives a 28 VDC signal command from EEC when all the following conditions are true: •
Single bleed air operation (from IASC)
•
Wing anti-ice is on (from IASC)
•
Bleed air is being extracted from the engine high-pressure port (from IASC)
•
Precooler outlet bleed temperature from the IASCs is above threshold for more than 10 seconds as determined by the EEC. The EEC determines the temperature limit based on altitude
•
Aircraft in flight
The L (R) ENG PCE DOOR OPEN advisory message is displayed on EICAS when the precooler exhaust (PCE) door is open. The EEC monitors the door latch actuator commands and reports failures to the onboard maintenance system (OMS). The PCE door opening function is also tested from the OMS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-26
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
Upper Piston Square Drive
Precooler Exhaust Door
Actuator Square Drive
Actuator Reset Check Hole
Striker Latch
Lower Piston Square Drive
EICAS OMS LEFT EEC DMC 1 and DMC 2
DMC EEC HP IASC WOW
Air/Gnd
Data Concentrator Unit Module Cabinet Electronic Engine Control High-Pressure Integrated Air System Controller Weight-On-Wheels ARINC 429 Discrete
Ch B Piston
IASC 1
Latch Actuator
NOTE Left engine PCE Door shown, right engine PCE Door similar.
CS1_CS3_3610_001
LEGEND
Ch A
Figure 13: Precooler Exhaust Door Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-27
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
BLEED AIR SOURCE SELECTION
•
Aircraft altitude is less than 16,000 ft
The IASC provides the logic for switching between the available bleed air sources.
•
Ice is detected above 1,500 ft or 2 minutes after takeoff
When the engines and the APU are not running, or their associated bleed switches are off, the IASC assumes a no bleed state. Pneumatic air from an external air source can then be used to supply air to the BAS. When bleed pressure is detected with no engine or APU source selected, the EXT AIR symbol appears on the AIR synoptic page. The IASC selects the APU as the default bleed air source, when the APU electronic control unit (ECU) provides a ready-to-bleed signal, and all of the following conditions apply: •
The APU BLEED PBAs is in the on (normal) position
•
The WING ANTI-ICE switch is in the AUTO position and the system has not been activated.
•
Slats and flaps extended (not fully retracted)
•
Aircraft altitude <15,500 ft
In addition, whenever both engines are selected OFF (or not running), as signaled from the EECs, the IASCs switch to the APU as the default bleed air source. When the IASC selects APU as the bleed air source, it signals the APU ECU to open the APU bleed air valve (BAV) and the cross-bleed valve (CBV). During and after engine start, the APU remains the default bleed air source. It supplies the environmental control system (ECS) on the ground and during takeoff to minimize the bleed demand on the engines. The IASC automatically switches to engine bleed air, provided either or both engines are running and the associated BLEED PBA is in the AUTO position, and any of the following conditions apply: •
APU BLEED PBA selected OFF or the APU is not ready to bleed
•
WING ANTI-ICE switch selected to ON or in AUTO and active
•
Slats and flaps fully retracted
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
When the IASC switches to the engine bleed air source, the APU BAV closes through the APU ECU and both PRSOVs are opened. During descent, with the APU available, the bleed air system switches from the engines to the APU when the aircraft altitude is less than 15,500 ft, flaps and slats are extended, and wing anti-ice is not required. The ENG BLEED MISCONFIG caution message is displayed on EICAS if both engine BLEED PBAs are selected OFF, and a low bleed air flow is detected from the APU. This indicates that the APU is unable to support the bleed air users while the APU is the bleed source. CROSS-BLEED VALVE OPERATION The cross-bleed valve (CBV) is directly controlled by the IASCs channel B. With the XBLEED switch in the AUTO (normal) position, the IASC automatically controls the opening and closing for the CBV. The crew can force the IASC to command the CBV open or closed by selection of the MAN OPEN or MAN CLSD positions. When the CBV is manually operated, XBLEED MAN OPEN or XBLEED MAN CLSD EICAS status messages are displayed. In the AUTO position, the IASC commands the CBV open whenever a single bleed source is available (one PRSOV open, APU bleed air source or external air). The CBV is closed when both PRSOVs are open. The ENG BLEED MISCONFIG caution message is displayed on EICAS if the XBLEED valve is selected to MAN OPEN while both engine BLEED PBAs are in the ON position. If the IASC detects that the CBV has failed to move (failed in position) when commanded, the XBLEED FAIL caution message is displayed on EICAS. This situation can be detected whether the XBLEED switch is selected to AUTO or any of the MAN positions. Under these conditions the IASC latches the EICAS message until an initiated built-in test (IBIT) is performed.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-28
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L PACK
OFF
XBLEED AUTO
MAN CLSD
FAIL
MAN OPEN
OFF
L BLEED FAIL OFF
R BLEED
Left EEC and Right EEC
SFECU 1 and SFECU 2
ADSPs
FAIL
APU BLEED
EICAS
OFF
FAIL
DMC 1 and DMC 2
OFF
OMS
ANTI-ICE WING AUTO
AIR PANEL
5 -40 PSI PSI
XBLEED
OFF
ON
5 -40 PSI
ANTI-ICE PANEL L PIPS 1 L PIPS 2
L PIPS 1 L PIPS 2
IASC 1 and IASC 2
ACS ADSP ATS BMPS BTS CBV CMD DMC ECU EEC FAV FTIS HPGC HPV IASC PIPS PRSOV SFECU WAIS M
APU
SYNOPTIC PAGE – AIR
LEFT ENGINE Fan
Users • L ACS • L WAIS • FTIS
T/M
FAV BMPS
PS
T/M
PRECOOLER
BTS
Fan Air Bleed Air Air Conditioning System Air Data System Probe Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Command Data Concentrator Unit Module Cabinet Electronic Control Unit Electronic Engine Control Fan Air Valve Fuel Tank Inerting System High-Pressure Ground Connection High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Slat/Flap Electronic Control Unit Wing Anti-Ice System ARINC 429 Motor
SOL
Solenoid
T/M
Torque Motor
Users • R ACS • R WAIS BTS
M
T/M
PRECOOLER
T/M
PS
BMPS
FAV
4th Stage
RIGHT ENGINE Fan 4th Stage
SOL
PRSOV
CBV
PRSOV
SOL
8th Stage
8th Stage HPV
To ATS
To ATS SOL
External Air HPGC
APU ECU
APU
HPV
CS1_CS3_3600_022
FAIL
LEGEND
R PACK
Figure 14: Bleed Air Source Selection and Cross-Bleed Valve Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
MONITORING AND TESTS The following page displays the crew alerting system (CAS) and INFO messages associated with the bleed air system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L2_BAS_MonitoringTests.fm
36 - Pneumatic 36-10 Bleed Air System
Table 3: STATUS Messages
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
MESSAGE
LOGIC
LOGIC
XBLEED MAN CLSD
Crossbleed valve manually selected and confirmed closed.
L BLEED FAIL
Left engine bleed failure or loss of left bleed leak detection or loss of both IASC 1 channels.
XBLEED MAN OPEN
Crossbleed valve manually selected and confirmed opened.
R BLEED FAIL
Right engine bleed failure or loss of right bleed leak detection or loss of both IASC 2 channels.
APU BLEED OFF
APU bleed selected OFF.
APU BLEED FAIL
The APU is not able to provide bleed air when requested or loss of bleed leak detection.
L BLEED OVHT
Left engine bleed overtemperature detected.
R BLEED OVHT
Right engine bleed overtemperature detected.
ENGINE BLEED MISCONFIG
XBLEED valve selected open and both engine bleed selected on or low flow on APU and left and right engine BLEED PBAs OFF.
XBLEED FAIL
XBLEED valve failed in position in AUTO or MAN mode.
Table 2: ADVISORY Messages MESSAGE AIR SYSTEM FAULT
Table 4: INFO Messages MESSAGE
LOGIC
36 L BLEED FAIL - L HPV FAIL CLSD
Left HPV detected failed closed by either left IASC channel.
36 L BLEED FAIL - L HPV FAIL OPEN
Left HPV detected failed open by either left IASC channel.
36 L BLEED FAIL - L PRESS REG SOV INOP
Left PRSOV causing a low or high-pressure.
36 L BLEED FAIL - L PRESS SOV FAIL OPEN
Left PRSOV detected failed open by either left IASC channel.
36 R BLEED FAIL - R HPV Right HPV detected failed closed by either right FAIL CLSD IASC channel.
LOGIC Loss of redundant or non-critical function for the air system (bleed or trim).
36 R BLEED FAIL - R HPV Right HPV detected failed open by either right FAIL OPEN IASC channel.
L ENG PCE DOOR OPEN L PCE door commanded opened (due to precooler overtemperature condition).
36 R BLEED FAIL - R PRESS REG SOV INOP
R ENG PCE DOOR OPEN R PCE door commanded opened (due to precooler overtemperature condition).
Right PRSOV detected failed open by either 36 R BLEED FAIL - R PRESS REG PRESS SOV right IASC channel. HPV FAIL OPEN
Table 3: STATUS Messages MESSAGE
LOGIC
L BLEED OFF
Left engine bleed selected OFF.
R BLEED OFF
Right engine bleed selected OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Right PRSOV causing a low or high-pressure.
36 L BLEED FAIL - L TEMP L BTS failure. SNSR INOP 36 R BLEED FAIL - R TEMP SNSR INOP
TECHNICAL TRAINING MANUAL
For Training Purposes Only
R BTS failure.
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36 - Pneumatic 36-10 Bleed Air System
Page Intentionally Left Blank
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36 - Pneumatic 36-10 Bleed Air System
Table 4: INFO Messages MESSAGE
LOGIC
36 AIR SYSTEM FAULT - L L BTS drift or single channel failure. BLEED TEMP SNSR REDUND LOSS 36 AIR SYSTEM FAULT R BLEED TEMP SNSR REDUND LOSS
R BTS drift or single channel failure.
36 AIR SYSTEM FAULT - L Left BMPS failure. BLEED MON PRESS SNSR INOP 36 AIR SYSTEM FAULT R BLEED MON PRESS SNSR INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Right BMPS failure.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L2_BAS_MonitoringTests.fm
36 - Pneumatic 36-10 Bleed Air System
PRACTICAL ASPECTS HIGH-PRESSURE VALVE, PRESSURE-REGULATING SHUTOFF VALVE, AND FAN AIR VALVE DEACTIVATION The high-pressure valve (HPV), pressure-regulating shutoff valve (PRSOV), and the fan air valve (FAV) can be secured closed for dispatch by rotating the manual override lever to the CLOSED position. The venting screw is used to secure the manual override lever in the closed position. Removing the vent screw from its stowed position vents the pneumatic actuator and prevents valve operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
Lanyard Venting Screw Insertion Point
Venting Screw HIGH-PRESSURE VALVE DEACTIVATED
CS1_CS3_3600_015
HIGH-PRESSURE VALVE
NOTE High-pressure valve shown, pressure-regulating shutoff valve and fan air valve similar.
Figure 15: High-Pressure Valve, Pressure-Regulating Shutoff Valve, and Fan Air Valve Deactivation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
CROSS-BLEED VALVE DEACTIVATION The cross-bleed valve (CBV) can be secured closed by rotating the manual override lever to the closed position and securing the lever with lockwire.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
Electrical Actuator
CS1_CS3_3600_023
Manual Override Lever
Figure 16: Cross-Bleed Valve Deactivation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
PRECOOLER EXHAUST DOOR MANUAL OPENING AND CLOSING The PCE door can be manually operated using a square drive tool. Open the PCE Door 1. Rotate the upper piston assembly square drive 90° counterclockwise. 2. Rotate the lower piston assembly square drive 90° counterclockwise.
Close the PCE Door 1. Rotate the upper piston assembly square drive 90° counterclockwise. 2. Rotate the lower piston assembly square drive 90° counterclockwise. 3. Rotate the striker latch square drive 60° counterclockwise.
3. Rotate the striker latch square drive 60° counterclockwise.
4. Rotate the actuator square drive 10.5° counterclockwise.
4. Manually pull the door to the open position. 5. Rotate the upper piston square drive 90° clockwise.
NOTE
6. Rotate the lower piston assembly square drive 90° clockwise. 7. Rotate the striker latch square drive 60° clockwise.
The square drive will return to initial position. 5. Manually push the pre-cooler exhaust door to the closed position. 6. Hold the pre-cooler exhaust door in the closed position and rotate the striker latch square drive 60° clockwise. 7. Insert a 1/8 in. allen key into the actuator reset check hole. Verify that the allen key can be inserted from 1.524 to 3.048 cm (0.6 to 1.2 in.). If not, repeat the previous steps. 8. Rotate the upper piston assembly square drive 90° clockwise. 9. Rotate the lower piston assembly square drive 90° clockwise.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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36 - Pneumatic 36-10 Bleed Air System
Upper Piston Square Drive Actuator Square Drive
Actuator Reset Check Hole
Striker Latch
CS1_CS3_3610_002
Lower Piston Square Drive
Figure 17: PCE Exhaust Door Manual Opening and Closing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
PCE DOOR TEST The PCE Door Test is used to detect any operational faults of the PCE system, including the electrical interface, solenoid, and door assembly.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
MAINT MENU EEC1-A-PCE DOOR TEST
1/4
*** WARNING *** PRE-COOLER EXT DOOR WILL OPEN DURING THIS TEST
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- FADEC IS POWERED
* NOTE * TEST WILL COMMENCE IMMEDIATELY, BUT THE DURATION IS UNTIL THE OPERATOR CONFIRMS DOOR OPERATION
PRESS TO START TEST
Write Maintenance Reports
Cancel
CS1_CS3_7321_025
PRESS THE START BUTTON BELOW TO INITIATE TEST
Figure 18: Precooler Exit Door Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-41
CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
ONBOARD MAINTENANCE SYSTEM TESTS On power-up, each channel of the IASCs performs a power-up built-in test (PBIT) to validate the input and system functions. During system operations each channel performs a continuous built-in test (CBIT) to validate input, active function operations, and proper active output operations. PBIT and CBIT monitor adjacent and opposite channel faults and failures, and store these in each channel’s memory for maintenance retrieval. Initiated built-in tests (IBIT) of the PRSOV, HPV, FAV, BTS and BMPS can be conducted on the OMS. Preconditions and instructions for the tests are provided on the test pages. The OMS also provides data pages for alarms status, valve position, and sensor status. Refer to ATA 45, Central Maintenance System for additional information.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
MAINT MENU
RETURN TO LRU/SYS OPS
IASC1 chA-IBIT BAS IBIT BAS 1/4 EQUIPMENT TESTED: PRSOV, HPV,FAV, BTS, BMPS PRECONDITIONS 123456-
MAKE MAKE MAKE MAKE MAKE MAKE
SURE SURE SURE SURE SURE SURE
WEIGHT ON WHEEL L BLEED SW IS SELECTED OFF R BLEED SW IS SELECTED OFF ALL ENGINE ARE OFF WING A/ICE SW IS SELECTED OFF APU BLEED SW IS SELECTED OFF TEST INHIBIT/INTERLOCK CONDITIONS
WEIGHT ON WHEEL LEFT BLEED SWITCH RIGHT BLEED SWITCH ENGINE STATUS WING A/ICE SWITCH APU BLEED SWITCH SWITCH
YYY YYY YYY YYYY YYY
Continue
Write Maintenance Reports
Cancel
CS1_CS3_3600_006
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Figure 19: Bleed Air System Initiated Built-In Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
36-21 BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM GENERAL DESCRIPTION The purpose of the bleed air leak and overheat detection system (BALODS) is to protect the aircraft from hot bleed air leaks, and subsequent overheat damage. The BALODS is a continuous monitoring system that detects the location of hot air leaks or overheat, and automatically isolates or reconfigures the affected system. The BALODS uses dual overheat detection loops that monitor the high-pressure ducting and air conditioning bays.
After an overheat detection and automatic reconfiguration, the flight crew may manually reconfigure the affected system from the AIR control panel. The BALODS system is active when the integrated air system controllers (IASCs) are powered. The BALODS function of the IASC is controlled by channel B, which is powered by DC ESS BUS 2 for IASC 1 and DC ESS BUS 3 for IASC 2.
Upon detection of overheat in the affected zone, the integrated air system controllers (IASCs) automatically command one or more valves to close within the following systems: •
Bleed air system: Left pressure-regulating shutoff valve (PRSOV), right PRSOV and cross-bleed valve (CBV)
•
Wing anti-ice system: Left and right wing anti-ice valves (WAIVs)
•
Air conditioning system: Left flow control valve (FCV) and right FCV
•
Trim air system (a sub-system of air conditioning): Left trim air shutoff valve (TASOV), and the right trim air pressure-regulating valve (TAPRV)
•
Auxiliary power unit (APU) bleed air: APU bleed air valve
Indication of the overheat condition is displayed on the engine indication and crew alerting system (EICAS) messages and on the AIR synoptic page. Failure of bleed leak detection capability is indicated by the EICAS messages and FAIL lights on the AIR panel. The IASCs are also capable of identifying and recording the location of the leak for use in troubleshooting. This information is available through the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
RAM AIR
BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM Overheat Detection Loops
BLEED AIR SYSTEM
L PRSOV
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
R PRSOV
L PACK
CBV
FAIL
DC ESS BUS 2 IASC 1
OFF
APU BLEED AIR VALVE
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL
APU ECU
OFF
OPEN
FAIL OFF R BLEED FAIL
APU BLEED
OFF
FAIL OFF
AIR PANEL Overheat Detection Loops
EICAS IASC 1 and IASC 2
DMC 1 and DMC 2
SYN OMS L WAIV R WAIV
L FCV
DC ESS BUS 3 IASC 2
AIR CONDITIONING SYSTEM
R FCV L TASOV R TAPRV
LEGEND CBV DMC ECU FCV IASC PRSOV TAPRV TASOV WAIV
Cross-Bleed Valve Data Concentrator Unit Module Cabinet Electronic Control Unit Flow Control Valve Integrated Air System Controller Pressure-Regulating Shutoff Valve Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Wing Anti-Ice Valve ARINC 429
CS1_CS3_3620_010
WING ANTI-ICE SYSTEM
Figure 20: BALODS Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-45
CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM OPERATION
Upon detection of overheat in the affected zone, the IASCs automatically perform the following actions within 15 seconds:
The dual-detection loops are installed along the high-pressure air ducts and associated components, in the following eight zones:
•
•
•
•
•
•
Left bleed manifold zone: Located downstream of the left engine pylon firewall and along the left inboard wing leading edge to the cross-bleed valve, the fuel tank inerting system (FTIS) heat exchanger, the auxiliary power unit check valve, the high-pressure ground connection, the trim air shutoff valve (TASOV), and on to the left flow control valve (FCV) Right bleed manifold zone: Located downstream of the right engine pylon firewall and along the right inboard wing leading edge to the cross-bleed valve (CBV), the trim air pressure-regulating valve (TAPRV), and on to the right FCV Left wing anti-ice zone: Located downstream of the left wing anti-ice valve to the left wing telescopic duct. There are no detection loops in the slats Right wing anti-ice zone: Located downstream of the right wing antiice valve to the right wing telescopic duct. There are no detection loops in the slats. Left pack zone: Located downstream of the left FCV to the mix manifold and associated aft cabin ducting
•
Right pack zone: Located downstream of the right FCV to the mix manifold and associated forward cabin ducting
•
Trim air duct zone: Located downstream of the left TASOV and right TAPRV to the respective ducting of the flight deck, forward cabin, aft cabin, and forward cargo trim air valves. In contrast to the forward and aft cabin, where only the trim air ducting is monitored, the complete air conditioning ducting to the flight deck, is monitored
•
APU duct zone: Located downstream of the APU bleed air valve to the APU check valve
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
•
Left bleed manifold zone: -
If on engine bleed, close the left pressure-regulating shutoff valve (PRSOV) and close the cross-bleed valve (CBV)
-
If on APU bleed, the IASCs sets the APU bleed request, to the APU electronic control unit (ECU) to false (closing the APU bleed air valve) and command the CBV to close
Right bleed manifold zone: -
If on engine bleed, close the right PRSOV and close the CBV
-
If on APU bleed, close the CBV
•
Left wing anti-ice zone: Close both wing anti-ice valves (WAIVs)
•
Right wing anti-ice zone: Close both WAIVs
•
Left pack zone: Close the left flow control valve (FCV)
•
Right pack zone: Close the right FCV
•
Trim air duct zone: Close the left TASOV and the right TAPRV
•
APU duct zone: Set the APU bleed request to false. The ECU, upon receiving the signal, closes the APU bleed air valve
The valves are locked out until the appropriate bleed air system (BAS) pushbutton annunciator (PBA) is selected OFF.
TECHNICAL TRAINING MANUAL
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36-46
CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Flight Compartment Left Aft Cabin
Left Fwd Cabin
Right Aft Cabin
Right Fwd Cabin MIX MANIFOLD FORWARD CARGO HEATING
Pressure Bulkhead LPGC
Pressure Bulkhead ERAV
TAV
TAV
TAV
TAV
LEFT PACK
RIGHT PACK FCV
LEFT ENGINE
Pylon Firewall
FCV
TASOV
Pylon Firewall
TAPRV
FTIS HEAT EXCHANGER
Precooler
RIGHT ENGINE Precooler
Telescopic Duct
CBV
WAIV
WAIV
Telescopic Duct
Tailcone Firewall
HPGC LEGEND FCV FTIS HPGC LPGC TAPRV TASOV TAV WAIV
Flow Control Valve Fuel Tank Inerting System High-Pressure Ground Connection Low-Pressure Ground Connection Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve Wing Anti-Ice Valve Check Valve
APU BLEED AIR VALVE
APU
CS1_CS3_3620_006
CBV ERAV
APU Duct Zone Left Bleed Manifold Zone Left Pack Zone Left Wing Anti-Ice Zone Right Bleed Manifold Zone Right Pack Zone Right Wing Anti-Ice Zone Trim Air Duct Zone Cross-Bleed Valve Emergency Ram Air Valve
Figure 21: BALODS Zones Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-47
CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
COMPONENT LOCATION The principal components of the bleed air leak and overheat detection system (BALODS) are: •
Integrated air system controllers
•
Dual-detection loops (Refer to figure 23)
•
Manifolds (Refer to figure 23)
INTEGRATED AIR SYSTEM CONTROLLERS The integrated air system controllers (IASCs) are located in the mid equipment bay forward rack.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
IASC 2
IASC 1
CS1_CS3_3615_018
MID EQUIPMENT BAY FWD RACK
INTEGRATED AIR SYSTEM CONTROLLERS (2X)
Figure 22: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-49
CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
DUAL-DETECTION LOOPS The dual-detection loops are located on high-pressure bleed air ducts and in areas of bleed air system components. MANIFOLDS The manifolds are located on the high-pressure air ducts and coupling covers.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Dual-Detection Loops
Manifold Coupling Cover
High-Pressure Air Duct
Coupling Cover Manifold
Manifold Wing Anti-Ice Valve
Manifold
Dual-Detection Loops
CS1_CS3_3615_012
LEFT WING ANTI-ICE DUCTING
NOTE Dual-detection loops and manifolds are similar in other areas of the aircraft high-pressure air ducts and components.
Figure 23: Bleed Air Leak and Overheat Detection System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
DETAILED COMPONENT INFORMATION DETECTION-LOOPS Dual-detection loops consist of parallel rows of sensing elements. The sensing elements are connected in series, either directly to the next element or using a wire to interconnect elements where there is little chance of exposure to high temperatures. The loops are identified as loop A and loop B. Both ends of each loop are connected to the integrated air system controller (IASC). The loops function independently to minimize false leak indications and improve dispatch reliability. The loops are powered by the integrated air system controller (IASC) with 15 VAC. When a loop is exposed to a hot air leak its resistance changes and is detected by the IASC. MANIFOLDS Dual-detection loops are installed on manifolds that are located on high-pressure air ducts and coupling covers. The purpose of the manifolds is to direct hot air leaks to the loops when a leak occurs. The loops are also routed over hot bleed air components in some zones to provide overheat damage protection. A hot bleed air leak in the duct assemblies is contained by the metal (stainless steel) shroud. Bleed air flows through the insulation air gaps and is routed to the manifold which directs the hot air onto the dual-detection loops, triggering an overheat condition. The manifolds on the coupling covers differ from those on the duct assemblies. A poppet valve (spring-loaded ball) is used to prevent nuisance warnings caused by the allowable natural leakage at the flange interfaces. This arrangement allows a minor buildup of pressure before the poppet valve opens releasing the hot air and heating the loops.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L3_BALODS_DetailedComponentInfo.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Dual-Detection Loops Detection Loop
Dual-Detection Loops
Spring
Ball
Detection Loop Coupling Cover
High-Pressure Air Duct COUPLING COVER MANIFOLD
MANIFOLD Shroud
Air Gap
Coupling Cover Manifold
Coupling Cover
Band Clamp (ref)
LEGEND Blead Air Leak
CROSS-SECTION OF HIGH-PRESSURE AIR DUCT ASSEMBLY
CS1_CS3_3615_013
V-Band Coupling (ref)
Figure 24: Manifold Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-53
CS130_36_L3_BALODS_DetailedComponentInfo.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
CONTROLS AND INDICATIONS The detection of a bleed air leak or overheat is displayed on the AIR synoptic page. Green flow lines indicate the normal presence of bleed air. When a leak is detected, the affected zone flow line(s) turn amber. A LEAK caution message is displayed on EICAS for the affected zone. White flow lines indicate that the particular zone has been confirmed isolated by the selection of the corresponding PBA or switch to OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-54
CS130_36_L2_BALODS_ControlsIndications.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
67$786
AIR
DOOR
ELEC
FLT CTRL
)8(/
HYD
AVIONIC
INFO
CB FLOW LINES
FWD
&.37 23 °C 22 °C
CARGO
15°C
Symbol
AFT
Condition
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
Valve displayed in white and closed position.
10 °C
15 °C
After isolation flow line turns to white no flow.
15 °C RAM AIR
TRIM AIR
Overheat displayed by amber flow line.
53$&.
/3$&.
40
XBLEED
36,
CAB ALT $38
ó3 LDG ELEV
5500 8.0 560
RATE CREW OXY
100 2000
SYNOPTIC PAGE – AIR
CS1_CS3_3620_007
40 36,
Figure 25: Bleed Air Leak and Overheat Detection System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-55
CS130_36_L2_BALODS_ControlsIndications.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
DETAILED DESCRIPTION BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM ACTIVATION Loop A of each dual-detection loop is connected to IASC 1 channel B, and loop B of each dual-detection loop is connected to IASC 2 channel B. Loop A and loop B of each detection loop is supplied with 15 VAC, and the IASC monitors the resistance of the current passing through each loop. During an overheat condition, the resistance of loop A and loop B changes. At a predetermined temperature, the IASC signals an overheat condition and automatically closes the valves in the affected system. Normal Mode Operation Under normal operation, detection by both of the dual-detection loops trigger the IASC with an overheat condition. The IASC automatically shuts down the affected zone and the appropriate crew alerting system (CAS) message is displayed on EICAS.
the affected zone to eliminate the risk of a bleed air leak. A LEAK DET FAIL caution message is displayed on EICAS for the affected zone and the associated FAIL light (L BLEED FAIL, R BLEED FAIL, L PACK FAIL or R PACK FAIL light) illuminates on the AIR panel. The crew are required to confirm the shutdown of the affected zone (s) by selecting the appropriate switches OFF. Selecting the switch to OFF resets the valve closing latch in the IASC. BALOD Failure If channel B of both IASCs fail, the entire BALODS is lost and considered failed and all zones are shut down. The LEAK DET FAIL caution message is displayed on EICAS. In this event the L and R BLEED pushbutton annunciators (PBAs) are selected OFF. If in flight, the aircraft must fly at a lower altitude as air conditioning is no longer available. Built-in Tests
Degraded Mode Operation
The IASC performs a power-up built-in test (PBIT), continuous BIT (CBIT), and landing BIT (LBIT). Initiated BIT (IBIT) can also be performed.
The BALODS operates in a degraded mode if one of the IASCs channel B is inoperative, or if loop A or loop B of the dual-detection loop fails. Loop failure is determined if one loop is shorted for more that 5 minutes, or one loop provides high-resistance (open) for more than 60 seconds.
PBIT is performed on every cold start on the ground. However, this cannot be done when bleed air is extracted as the system performs a calibration test of the loops and recording the resistance of the detection loops. This resistance value is stored by the IASC for use in detecting the location of bleed air leaks.
The LEAK DET FAULT advisory message is displayed on EICAS, indicating the fault and lost redundancy.
CBIT continually checks the loops for shorts and open circuits (continuity). A failure is used to switch to the degraded mode. The failure is indicated by the LEAK DET FAULT advisory message.
In the degraded mode, overheat detection relies on the remaining single loop to trigger the overheat annunciation for the affected zone. Failure Mode Operation In the event of a dual-loop failure in a particular zone, detection capabilities for that zone are lost. The IASC automatically shuts down Copyright © Bombardier Inc. CS130-21.07-06-00-121418
While the PBIT does a complete system test (shorts, continuity, etc) and calibration, the L-BIT only checks for loop continuity and internal system 5 minutes after landing. A loop failure is indicated by the LEAK DET FAULT advisory message.
TECHNICAL TRAINING MANUAL
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36-56
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK
R PACK XBLEED
FAIL
MAN CLSD
OFF
AUTO
MAN OPEN
L BLEED FAIL OFF
OPEN
EICAS
FAIL
DMC 1 and DMC 2
OFF R BLEED
SYN
FAIL
APU BLEED
OMS
OFF
FAIL OFF
AIR PANEL
Typical Zone Detection Loops
ANTI-ICE WING AUTO
OFF
15 VAC
ON
Loop A 15 VAC Loop B
CH A
CH B
CH B
IASC 1
LEGEND IASC
Integrated Air System Controller ARINC 429
CH A
IASC 2
DC ESS BUS 2
DC ESS BUS 3
CS1_CS3_3615_014
ANTI-ICE PANEL
Figure 26: Bleed Air Leak and Overheat Detection System Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-57
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
BALODS LEAK LOCALIZATION The BALODS performs leak localization using software embedded in the IASCs. Leak location accuracy is within 50.8 cm (20 in.). The operating principle is based on the linear electrical resistance of the sensing elements inner conductor. Upon leak detection, the increasing temperature decreases the resistance between the inner and outer conductor. Software in the IASC processes this change in resistance with total loop resistance without a leak, and uses a ratio of the two resistances to indicate a leak location. Calibration of the loop resistance is performed automatically on power-up except if power-up occurs during bleed air extraction. BALODS leak localization by maintenance personnel is obtained through the OMS and compared to a reference table in ATA 05-51-44 of the aircraft maintenance publication (AMP).
NOTE The bleed air extraction system MUST BE OFF during the LEAK LOCALIZATION test to prevent the radiant heat from the bleed extraction from affecting the resistance value of the loops.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-58
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
LEGEND Data Concentrator Unit Module Cabinet Integrated Air System Controller ARINC 429
IASC1 CH A - Sensor Status
Write Maintenance Reports ZONE
EICAS DMC 1 and DMC 2
SYN OMS
Typical Zone Dual-detection Loops 2
In
Loop A 15 VAC
CH B
CH A
Out 1 1 CH A
CH B
IASC 1
Loop B 15 VAC
In
2
Out IASC 2
NOTE 1
Loop A and Loop B shorted by exposure to hot air leak, and triggers a change in resistance.
2
Software in IASC uses a ratio of each loop total resistance without leak, and total resistance with leak, to determine the location of the leak.
RETURN TO FAULT MSGS
IASC1 CHB
IASC2 CHB
001 / 002
L L L L
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
R R R R
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
L L L L
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
R R R R
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
L L L L
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
R R R R
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
% CS1_CS3_3615_016
DMC IASC
MAINT MENU
Figure 27: BALODS Leak Localization Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-59
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the bleed air leak and overheat detection system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-60
CS130_36_L2_BALODS_MonitoringTests.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
CAS MESSAGES Table 7: INFO Message
Table 5: CAUTION Messages MESSAGE
LOGIC
MESSAGE
LOGIC
36 LEAK DET FAULT - Loss of one detection loop in the wing anti-ice zone. LOOP REDUND LOSS
L BLEED LEAK
Left bleed leak detected.
R BLEED LEAK
Right bleed leak detected.
APU BLEED LEAK
APU bleed leak detected.
L PACK LEAK
Left pack leak detected.
R PACK LEAK
Right pack leak detected.
TRIM AIR LEAK
Trim air leak detected.
WING A/ICE LEAK
Wing anti-ice leak detected.
LEAK DET FAIL
Entire leak detection system failed and either L BLEED, R BLEED, or APU BLEED PBA selected ON.
Table 6: ADVISORY Message MESSAGE LEAK DET FAULT
LOGIC One leak detection loop failed. Refer to INFO messages.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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36-61
CS130_36_L2_BALODS_MonitoringTests.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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36-62
CS130_36_ILB.fm
ATA 47 - Liquid Nitrogen (Fuel Tank Inerting System)
BD-500-1A10 BD-500-1A11
47 - Liquid Nitrogen (Fuel Tank Inerting System)
Table of Contents 47-10 Fuel Tank Inerting System Generation.........................47-2
47-20 Fuel Tank Inerting System Distribution...................... 47-36
General Description .........................................................47-2 Component Location .......................................................47-4 Component Information ...................................................47-6 Heat Exchanger ...........................................................47-6 Air Separation Module ................................................47-8 Detailed Description ......................................................47-10
General Description ...................................................... 47-36 Component Location ..................................................... 47-38 Check Valves ............................................................ 47-38 Flame Arrestor .......................................................... 47-38 Flow Balancing Orifices............................................. 47-38 Maintenance Port Plugs ............................................ 47-38
47-00 Fuel Tank Inerting System Control and Indicating...........................................................47-12 General Description .......................................................47-12 Component Location .....................................................47-14 Inerting Control Unit ..................................................47-16 Component Information .................................................47-18 Inlet Isolation Valve....................................................47-18 Temperature Isolation Valve ......................................47-18 Dual-Flow Shutoff Valve ............................................47-18 Detailed Description ......................................................47-20 Fuel Tank Inerting System Interface ..........................47-20 Inerting Control Unit ...................................................47-22 Normal Operating Mode.............................................47-24 Health Monitoring Mode.............................................47-28 Inhibit Mode ...............................................................47-28 Latched Shutdown Mode ...........................................47-28 Maintenance Mode ....................................................47-28 Monitoring and Test .......................................................47-30 CAS Messages ..........................................................47-31 Practical Aspects ............................................................47-32 Inlet Isolation Valve Lockout (TBC)............................47-32 Onboard Maintenance System ..................................47-34
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-i
CS130_M5_47_NitrogenTOC.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System)
List of Figures Figure 1: Fuel Tank Inerting System Generation..................47-3 Figure 2: Fuel Tank Inerting System Generation Components .......................................47-5 Figure 3: Heat Exchanger Assembly ....................................47-7 Figure 4: Air Separation Module...........................................47-9 Figure 5: Fuel Tank Inerting System Generation Detailed Description ............................................47-11 Figure 6: Fuel Tank Inerting System Control and Indicating .........................................47-13 Figure 7: Fuel Tank Inerting System Control Components ...........................................47-15 Figure 8: Inerting Control Unit ............................................47-17 Figure 9: Fuel Tank Inerting System Control Components ...........................................47-19 Figure 10: Fuel Tank Inerting System Interface ...................47-21 Figure 11: Inerting Control Unit ............................................47-23 Figure 12: Fuel Tank Inerting System Start Sequence.........47-25 Figure 13: Flow Control, Overtemperature , and Overpressure Protection ............................47-27 Figure 14: Fuel Tank Inerting System Modes.......................47-29 Figure 15: Inlet Isolation Valve Lockout................................47-33 Figure 16: Onboard Maintenance System............................47-35 Figure 17: Fuel Tank Inerting System Distribution ...............47-37 Figure 18: Fuel Tank Inerting System Distribution Components .....................................47-39
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47-ii
CS130_M5_47_NitrogenLOF.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System)
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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47-iii
CS130_M5_47_NitrogenLOF.fm
LIQUID NITROGEN - CHAPTER BREAKDOWN (FUEL TANK INERTING SYSTEM)
LIQUID NITROGEN (FUEL TANK INERTING SYSTEM) CHAPTER BREAKDOWN
Fuel Tank Inerting System Generation
1 Fuel Tank Inerting System Distribution
2 Fuel Tank Inerting System Indication
3 Fuel Tank Inerting System Control
4
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
47-10 FUEL TANK INERTING SYSTEM GENERATION GENERAL DESCRIPTION The fuel tank inerting system (FTIS) includes an onboard inert gas generation system (OBIGGS). The OBIGGS uses engine bleed air, auxiliary power unit (APU) air, or air from a ground cart to supply nitrogen-enriched air (NEA) to the fuel tanks. The inert NEA floods the empty space above the fuel, thus reducing the flammability in the fuel tanks. The system is controlled automatically through the inerting control unit (ICU). The bleed air is routed to the ozone converter, through an inlet isolation valve (IIV), to reduce the ozone concentration. The air supplied by the ozone converter passes through the heat exchanger assembly where it is cooled by the ram air. Some bleed air bypasses the heat exchanger via the temperature control valve (TCV). The TCV modulates to maintain the heat exchanger outlet temperature within a range that provides optimal OBIGGS performance. A ground cooling fan provides cooling air flow through the heat exchanger when the aircraft is at low airspeeds. A ground cooling fan check valve closes when the fan is operating. The check valve ensures that air is drawn over the heat exchanger during fan operation. The cooled air is filtered by the air separation module (ASM) inlet filter. Water vapor from the supply airflow is removed through a coalescer within the filter. The water drains from the filter assembly into the aircraft belly through a drain orifice in the bottom of the filter bowl. The cooled and filtered bleed air is supplied to the air separation module (ASM), which removes the oxygen from the air and creates the NEA.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-2
CS130_47_L2_FTISG_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
DC BUS 1
Ram Air Inlet Temperature Control Valve
ICU
S/M
SOL
O2P
P
OZONE CONVERTER
SOL
AIR SEPARATION MODULE
Bleed Air IIV
SOL SOL
To Fuel Tanks
ASM INLET FILTER
HEAT EXCHANGER ASSEMBLY
Ground Cooling Fan
HEAT EXCHANGER COOLING CHECK VALVE
LEGEND
Ground Cooling Fan Run Signal to ICU Ram Air Exhaust
ICU IIV
Bleed Air Cooling Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Inerting Control Unit Inlet Isolation Valve
CS1_CS3_4710_008
DC BUS 1
Figure 1: Fuel Tank Inerting System Generation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-3
CS130_47_L2_FTISG_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
COMPONENT LOCATION The following components of the FTIS are installed in the wing-to-body fairing (WTBF): •
Heat exchanger
•
Ground cooling fan
•
Heat exchanger cooling check valve
•
Ozone converter
•
Air separation module inlet filter
•
Air separation module
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-4 CS130_47_L2_FTISG_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
Air Separation Module
Air Separation Module Inlet Filter Ozone Converter Heat Exchanger Cooling Check Valve Heat Exchanger
CS1_CS3_4710_010
Ground Cooling Fan
Figure 2: Fuel Tank Inerting System Generation Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-5 CS130_47_L2_FTISG_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
COMPONENT INFORMATION HEAT EXCHANGER The heat exchanger consists of a cross flow heat exchanger element, a temperature control valve (TCV), and associated bypass ducting. The heat exchanger regulates the temperature of the air entering the air separation module (ASM), by mixing hot bleed air with air cooled by the heat exchanger. The TCV is a butterfly valve driven by a stepper motor. The ICU commands the TCV based on temperature signals from the ASM inlet temperature sensor. Limit switches on the valve provide position indication to the ICU. If the stepper motor fails, the valve remains in the last position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-6 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
Cooled Air Out Ram Air Inlet
Bleed Air In Ram Air Outlet
CS1_CS3_4710_004
Temperature Control Valve
Figure 3: Heat Exchanger Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-7 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
AIR SEPARATION MODULE The air separation module (ASM) is constructed of thousands of hollow fiber membranes that run the length of the module. The ASM separates the nitrogen and oxygen components of the bleed air into individual gas streams, using a process known as selective permeation. Air contains 78% nitrogen, 21% oxygen, and 1% other gases. Each gas has a characteristic permeation rate that is a function of its ability to dissolve and diffuse through a membrane. This characteristic allows fast gases, such as oxygen, to be separated from slow gases, such as nitrogen. Bleed air enters the ASM inlet at a controlled pressure and temperature. The bleed air feeds through hollow fiber membranes. The oxygen permeates through the walls of the membrane and discharges through the oxygen-enriched air (OEA) outlet. The remaining nitrogen-enriched air (NEA) flows down the hollow fiber membranes, to the NEA outlet, and on to the fuel tanks.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-8 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
Nitrogen-Enriched Air Outlet
Oxygen-Enriched Air Outlet
Flow
Hollow Fiber
CS1_CS3_4710_009
Inlet
Figure 4: Air Separation Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-9 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
DETAILED DESCRIPTION If the amount of ram cooling air is insufficient, the ground cooling fan (GCF) provides the cooling air flow to the heat exchanger assembly. The ICU uses the aircraft mach number information from the data concentrator unit module cabinet (DMC) to determine when the GCF should be running. The GCF runs anytime the mach number is less than 0.2 Mach. The GCF is powered by a solid-state power controller (SSPC), therefore a turn ON request is sent to the DMCs to activate the SSPC. At startup, the ICU energizes the GCF before opening the inlet isolation valve (IIV). This ensures that the heat exchanger is capable of cooling the initial flow of bleed air to avoid nuisance fuel tank inerting system (FTIS) shutdowns. The GCF sends a run signal to the ICU to indicate that it is operating. The fan speed and overspeed protection is managed by the fan control software. The fan automatically shuts down when overspeed is detected and a fault signal is set. When the ICU has determined the bleed air supply is within the normal range, the IIV opens to allow bleed air flow into the FTIS. The temperature control valve (TCV) is modulated to maintain the proper air separation module (ASM) inlet temperature of 60°C (140°F). If the heat exchanger outlet temperature exceeds 60°C (185°F), the TCV is driven towards the closed position and is fully closed at 85°C (185°F).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-10 CS130_47_L3_FTISG_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
DMC 1 AND DMC 2
ICU
CDC 1-11-7 Fuel Inert CTLR DC SSPC 5A BUS 1
Ram Air Inlet Temperature Control Valve
Logic: Always On
S/M
OZONE CONVERTER
SOL
TIV
Logic: Single Input
SOL SOL
To Fuel Tanks DFSOV
ASM INLET FILTER
HEAT EXCHANGER ASSEMBLY CDC 1-10-1 Electric Ground Cooling Fan DC BUS 1 SSPC 15A
O2P AIR SEPARATION MODULE
Bleed Air IIV
OXYGEN SENSOR
PRESSURE SENSOR ASM INLET TEMPERATURE P SENSOR
Ground Cooling Fan
HEAT EXCHANGER COOLING CHECK VALVE
LEGEND
Ground Cooling Fan Run Signal to ICU Ram Air Exhaust
DFSOV ICU IIV TIV
Bleed Air Cooling Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Dual-Flow Shutoff Valve Inerting Control Unit Inlet Isolation Valve Temperature Isolation Valve ARINC 429
CS1_CS3_4710_005
SOL
HEAT EXCHANGER OUTLET TEMPERATURE SENSOR
Figure 5: Fuel Tank Inerting System Generation Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-11 CS130_47_L3_FTISG_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
47-00 FUEL TANK INERTING SYSTEM CONTROL AND INDICATING GENERAL DESCRIPTION The inerting control unit (ICU) receives power from DC BUS 1. The ICU powers all of the fuel tank inerting system (FTIS) electrical components, except the ground cooling fan (GCF), which also receives power from DC BUS 1. The ICU software is field loadable. The FTIS enters the normal operating mode when the aircraft is fully powered, bleed air is available, and the ICU power-up built-in test (BIT) is complete. Operation of the FTIS is fully automatic. The inlet isolation valve (IIV) is the primary ON/OFF control for the FTIS. The IIV isolates the FTIS from the pneumatic system. The valve moves to the closed position for a loss of power or pneumatic pressure. The temperature isolation valve (TIV), located at the inlet to the air separation module (ASM), is used to shut off air flow to the distribution system. This prevents hot air from contacting fuel or fuel vapor during an overtemperature condition. When an overtemperature is detected, the valve is commanded closed. The TIV is an electrically-controlled, pneumatically-operated gate valve. A microswitch provides position sensing to the ICU. The dual-flow shutoff valve (DFSOV) supplies the distribution system with nitrogen-enriched air (NEA) at a flow rate based on altitude and vertical speed. It also prevents fuel and fuel vapor migration to the FTIS when the system is not operating. The DFSOV outputs three flow rates. The low-flow solenoid is energized on the ground, and in climb and cruise. The medium-flow solenoid is used in approach and slow descent, and the high-flow rate is used in descent. The high-flow is created by energizing both solenoids. System monitoring is limited to crew alerting system (CAS) messages indicating a fault has occurred and the system is shut down. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Temperature and pressure sensors are used for control, system health monitoring, and provide overtemperature and overpressure shutdown protection. The heat exchanger outlet temperature sensor monitors the heat exchanger discharge temperature. The air separation module (ASM) inlet temperature sensor monitors the temperature of the bleed air entering the ASM. The oxygen sensor monitors the outlet of the ASM. The sampling port is located at the inlet to the dual-flow shutoff valve. If the oxygen level concentration is detected above the preset threshold, a crew alerting system (CAS) INFO message is displayed. System monitoring is limited to CAS messages that indicate a fault has occurred and the system is degraded or shut down. The FTIS has five operating modes: •
Normal operating mode - the FTIS provides NEA when bleed air is available
•
Health monitoring mode - the ICU performs a pressure and oxygen concentration data check of the NEA once during cruise
•
Inhibit mode - power is removed from all FTIS components. The system recovers to the normal operating mode when the inhibit condition no longer exists
•
Latched shutdown mode - an overtemperature, overpressure fault, or sensor fault is detected and the system shuts down
•
Maintenance mode - provides the capability to test and troubleshoot FTIS faults to the line replaceable unit (LRU) level
TECHNICAL TRAINING MANUAL
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47-12 CS130_47_L2_FTISCI_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
DC BUS 1
ICU
Ram Air Inlet Temperature Control Valve
S/M
SOL
O2P
P
OZONE CONVERTER
SOL
Bleed Air IIV
TIV HEAT EXCHANGER ASSEMBLY
Heat Exchanger Outlet Temperature Sensor
SOL SOL
AIR SEPARATION MODULE ASM Inlet Temperature Sensor
To Fuel Tanks DFSOV
Oxygen Sensor
Pressure Sensor
Ground Cooling Fan
HEAT EXCHANGER COOLING CHECK VALVE
LEGEND
Ground Cooling Fan Run Signal to ICU ASM DFSOV ICU IIV TIV
Ram Air Exhaust
Bleed Air Cooling Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Air Separation Module Dual-Flow Shutoff Valve Inerting Control Unit Inlet Isolation Valve Temperature Isolation Valve
CS1_CS3_4730_002
DC BUS 1
Figure 6: Fuel Tank Inerting System Control and Indicating Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-13 CS130_47_L2_FTISCI_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
COMPONENT LOCATION The following fuel tank inerting control and indicating system components are located in the wing-to-body fairing (WTBF): •
Inlet isolation valve
•
Temperature isolation valve
•
Dual-flow shutoff valve
•
Heat exchanger outlet temperature sensor
•
Pressure sensor
•
Air separation module (ASM) inlet temperature sensor
•
Oxygen sensor
•
Inerting control unit (refer to figure 8)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-14 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Dual-Flow Shutoff Valve
Oxygen Sensor Temperature Isolation Valve
Inlet Isolation Valve
Pressure Sensor
ASM Inlet Temperature Sensor
CS1_CS3_4730_010
Heat Exchanger Outlet Temperature Sensor
Figure 7: Fuel Tank Inerting System Control Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-15 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
INERTING CONTROL UNIT The inerting control unit (ICU) is located in the mid equipment bay.
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47-16 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
CS1_CS3_4730_008
MID EQUIPMENT BAY AFT RACK
INERTING CONTROL UNIT
Figure 8: Inerting Control Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-17 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
COMPONENT INFORMATION INLET ISOLATION VALVE The inlet isolation valve (IIV) is a solenoid-controlled, pneumatically-operated butterfly valve. The valve has a position indicator that must not be used to move the valve. TEMPERATURE ISOLATION VALVE The temperature isolation valve (TIV) is an electrically-controlled, pneumatically-operated gate valve. A microswitch provides position sensing to the ICU. The TIV also has a position indicator. DUAL-FLOW SHUTOFF VALVE The DFSOV is a dual-poppet, 28 VDC electrically-controlled, pneumatically-operated valve. The poppets are spring-loaded closed when the solenoids are de-energized. Visual indicators show when the poppets are in the open or closed position. Microswitches provide position information to the ICU when the poppets are closed.
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47-18 CS130_47_L2_FTISCI_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Inlet
Position Indicator
Position Indicator
Valve Position Indicator
Outlet
INLET ISOLATION VALVE Position Indicator TEMPERATURE ISOLATION VALVE
CS1_CS3_4730_003
DUAL-FLOW SHUTOFF VALVE
Figure 9: Fuel Tank Inerting System Control Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-19 CS130_47_L2_FTISCI_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
DETAILED DESCRIPTION FUEL TANK INERTING SYSTEM INTERFACE Communication with the aircraft systems is through the data concentrator unit module cabinet (DMC). The ICU communicates with DMC 1. If DMC 1 becomes unavailable, the ICU will switch to DMC 2. The ICU transmits system and component health information, and receives flight phase and conditions information that is used by the ICU to determine the necessary nitrogen-enriched air (NEA) flow. The ARINC 429 inputs received via the DMC are from: •
Air data system for environmental data and Mach number
•
Fuel quantity computer for refuel status
•
Integrated air system controller (IASC) for bleed pressure
•
Onboard maintenance system (OMS) for testing
The digital channel controls the ground for the IIV and the TIV. Each analog circuit controls the 28 VDC for its respective valve. Analog channel 1 controls the power to the IIV, and analog channel 2 controls power to the TIV. In normal operation, the digital channel will remove the grounds from the IIV, TIV, and DFSOV when an overtemperature is sensed. If the digital channel fails, and the heat exchanger outlet temperature sensor senses an overheat, analog channel 1 removes power from the IIV. If the ASM inlet temperature sensor detects an overheat, analog channel 2 removes power from the TIV. The temperature control valve (TCV) position is based on both the heat exchanger outlet temperature sensor, and the ASM inlet temperature sensor. The ASM inlet sensor is the primary driver of temperature, but the heat exchanger outlet sensor provides a safety function and commands the valve to full-closed at 85°C (185°F).
The ICU remains in inhibit mode while performing power-up BIT (PBIT). After completing a successful PBIT, the ICU begins energizing valves in a controlled startup sequence. As each valve is energized, a position sensor check is performed and other safety related parameters are checked. The ICU digital channel is the primary monitor of FTIS temperature. Two independent temperature monitoring analog backup channels are provided for shutdown of the FTIS in case of a digital channel failure.
The pressure sensor at the air separation module (ASM) inlet is used for health monitoring and detection of overpressure conditions that could damage system components. The oxygen sensor provides an indication of the oxygen concentration and absolute pressure of the nitrogen-enriched air (NEA) delivered by the inerting system. The output from the oxygen sensor is used to monitor the performance of the ASM.
The heat exchanger outlet temperature sensor monitors the heat exchanger discharge temperature. The sensor is a dual-element type, One element reports to analog channel 1 of the inerting control unit (ICU), and the other element reports to the digital channel. The ASM inlet temperature sensor monitors the air temperature entering the ASM. This temperature is used by the inerting control unit (ICU) for temperature control and health monitoring. It is a dual-element type. One element reports to analog channel 2 of the ICU, and the other element reports to the digital channel. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-20 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
CDC 1-10-1 Electrical Ground Cooling Fan DC BUS 1 SSPC 15A
GROUND COOLING FAN
INERTING CONTROL UNIT
Logic: Single Input
Digital Channel GCF Run Signal
CDC 1-11-7 Fuel Tank Inerting System DC BUS 1 SSPC 5A
Position
INLET ISOLATION VALVE
Ground
Logic: Always On
Analog Channel 1 28 VDC HEAT EXCHANGER OUTLET TEMPERATURE SENSOR
Temperature
Temperature 28 VDC Fuel Quantity Computer
TEMPERATURE CONTROL VALVE
Position Position
TEMPERATURE ISOLATION VALVE
Ground EICAS
28 VDC Pressure DMC 1 and 2
OMS
PRESSURE SENSOR
Analog Channel 2 28 VDC
ASM INLET TEMPERATURE SENSOR
Temperature Air Data
IASC
LEGEND ASM GCF IASC
Air Separation Module Ground Cooling Fan Integrated Air System Controller ARINC 429
Temperature O2 Signal 28 VDC 28 VDC Position
OXYGEN SENSOR DUAL-FLOW SHUTOFF VALVE
CS1_CS3_4700_006
HMU
Figure 10: Fuel Tank Inerting System Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-21 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
INERTING CONTROL UNIT The inerting control unit (ICU) has a digital channel that is the primary monitor of FTIS temperatures. Two independent temperature monitoring analog channels shut down the FTIS in case of failure of the digital channel. When the digital or analog monitor circuits detect a failure, the controller enters a latched shutdown mode. The ICU provides automatic control of the FTIS. Primary functions include: •
System startup
•
System control
•
Temperature control
•
Overtemperature protection
•
Overpressure protection
•
Health monitoring
•
System shutdown
•
System inhibits
Secondary functions are part of a maintenance mode, which includes activating and monitoring each FTIS component as commanded by the aircraft's maintenance computer.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-22 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
System Startup System Control Analog Channel 1
Temperature Control Overtemperature Protection Overpressure Protection
Digital Channel
Health Monitoring
Analog Channel 2
System Shutdown System Inhibits Maintenance Modes
CS1_CS3_4740_005
INERTING CONTROL UNIT
Figure 11: Inerting Control Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-23 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
NORMAL OPERATING MODE The normal operating mode consists of the following submodes: •
Starting sequence
•
Flow control
•
Overpressure protection
•
Temperature control and overtemperature protection
6. Check that the heat exchanger outlet temperature sensor does not indicate overtemperature. 7. Open the TIV. 8. Check that ASM inlet temperature sensor does not indicate overtemperature.
When initially energized with 28 VDC, the ICU remains in inhibit mode while performing power-up BIT (PBIT). After completing a successful PBIT, the FTIS ICU begins energizing valves in a controlled startup sequence. As each valve is energized, a position sensor check is performed and other safety related parameters are checked.
9. Open mid flow solenoid of the dual-flow shutoff valve (DFSOV) for 30 second purge and then open low flow solenoid of the DFSOV. 10. If bleed air is not available, enter the inhibit mode. 11. If bleed air is available, enter the operating mode and begin temperature and flow control.
If the amount of ram cooling air is insufficient, the ground cooling fan (GCF) provides the cooling air flow to the heat exchanger assembly. The GCF runs if the aircraft mach number is less than 0.2 M. When the ICU is satisfied with the initialization checks, and has determined the bleed air supply is within the normal range, the IIV opens to allow bleed air flow into the FTIS. The temperature control valve (TCV) is modulated to maintain the proper air separation module (ASM) inlet temperature and the FTIS is considered powered-up and operational.
NOTE The ICU uses a bleed pressure signal from the IASC to determine if there is enough bleed air available to open the IIV and start the fuel tank inerting process.The PBIT test runs even if bleed air is not available. Since the IIV, TIV, and DFSOV do not open if bleed air is not available, they are only monitored for the closed position. The bleed pressure signal is also used to determine if the IIV has failed.
Starting Sequence The ICU controls the system startup as follows: 1. Perform a power-up BIT (PBIT). 2. Begin temperature monitoring. 3. Command the GCF to run if the aircraft is on ground, and check for GCF feedback. 4. Cycle the TCV open and closed. Confirm valve closed. 5. Open the inlet isolation valve (IIV) and confirm valve position. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-24 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Power-up BIT
Temperature monitoring
Ground cooling fan Ground cooling fan run signal
Temperature control valve cycled open Temperature control valve closed
Inlet isolation valve open Inlet isolation valve position feedback INERTING CONTROL UNIT Heat exchanger outlet temperature sensor detects no overtemperature
Open temperature isolation valve Temperature isolation valve position feedback
Dual-flow shutoff valve mid-solenoid open
No bleed air FTIS enters inhibit mode
30 second purge Low solenoid open
Bleed air available FTIS enters normal operating mode
CS1_CS3_4730_009
Air separation module inlet temperature sensor detects no overtemperature
Figure 12: Fuel Tank Inerting System Start Sequence Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-25 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Flow Control
Temperature Control and Overtemperature Protection
The ICU energizes the appropriate dual-flow shutoff valve (DFSOV) solenoids based on aircraft altitude and vertical speed.
The ICU maintains the ASM inlet temperature at 60°C (140°F). The ICU uses ASM inlet temperature sensor and HE outlet temperature sensor information to modulate the temperature control valve (TCV).
When low-flow is required, only the low solenoid is energized. For mid-flow, the mid solenoid is energized. High-flow is achieved with both solenoids energized. During ground, climb, and cruise operations, the fuel tank inerting system (FTIS) operates in low-flow mode. During this time, the NEA reduces the fuel tank oxygen concentration and fills the fuel tank space, storing NEA to minimize the amount required to be generated during descent. As the aircraft climbs, the NEA will be venting overboard in order to maintain a pressure slightly above ambient in the tank.
An overtemperature sensed by the digital channel closes all of the valves. An overtemperature sensed by either of the analog channels results in its respective valve closing. The controller enters the latched shutdown mode when either the digital or analog monitor circuits detect an overtemperature. The temperature shutdown thresholds are as follows: Digital:
During descent, a vent inflow condition occurs to maintain ambient pressure. The FTIS enters the high-flow mode to ensure that there is adequate mixing of the NEA and vent air.
•
85°C (185°F) for 2 minutes
•
88°C (190°F)
Overpressure Protection
•
Analog: 90°C (194°F)
The ICU provides overpressure protection of the FTIS components and ducting. If an overpressure condition of 60 psi is detected at the inlet to the air separation module ASM, the FTIS is automatically shut down by closing the inlet isolation valve (IIV) and DFSOV. The temperature inlet valve (TIV) is held open for a short period of time to allow any trapped pressure to bleed off through the filter drain and ASM oxygen-enriched air (OEA) port and then closed. Overpressure protection of the fuel tank is provided by the fixed orifices in the DFSOV.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-26 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
HE OUTLET TEMPERATURE SENSOR B
ASM INLET TEMPERATURE SENSOR B
Digital Channel
DFSOV Ground Cruise Climb
Low-Flow Solenoid
ICU
Fast Descent IIV Mid-Flow Solenoid
Slow Descent
SOL
TIV
SOL
DFSOV
SOL SOL
ICU Analog Channel 1
Analog Channel 2
Flow Control HE OUTLET TEMPERATURE SENSOR A
ASM INLET TEMPERATURE SENSOR A
Temperature Control and Overtemperature Protection
LEGEND
ICU Overpressure Sensor
SOL
SOL
IIV
TIV
ASM INLET FILTER
PS SOL SOL
ASM DFSOV
Overpressure Protection
OEA Port Trapped Pressure
Trapped Pressure
CS1_CS3_4740_006
ASM DFSOV HE ICU IIV OEA TIV
Bleed Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Air Separation Module Dual-Flow Shutoff Valve Heat Exchanger Inerting Control Unit Inlet Isolation Valve Oxygen Enriched Air Temperature Isolation Valve
Figure 13: Flow Control, Overtemperature, and Overpressure Protection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-27 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
HEALTH MONITORING MODE Although oxygen sampling is done continuously, the FTIS enters the health monitoring mode one time per flight. This occurs during the operating mode when the aircraft is in the cruise flight phase and above 18,000 ft. When the health monitoring interval is initiated, the oxygen sensor will be turned on to measure the NEA oxygen concentration at the ASM outlet for a duration of 1 minute. A small gas sample, drawn from the outlet of the ASM, is fed to the oxygen sensor. The pressure and oxygen concentration is sent to the ICU, where it is used in conjunction with other FTIS and aircraft parameters to determine system health. An INFO message is displayed on the engine indication and crew alerting system (EICAS) when the oxygen concentration is out of limits over a period of several flights.
The ICU exits inhibit mode and starts normal FTIS operation when the conditions causing the inhibit mode are no longer present for at least 3 seconds. LATCHED SHUTDOWN MODE In the latched shutdown mode, all FTIS components de-energize with no automatic recovery. The ICU enters into latched shutdown mode if one of the following events occurs: •
Overtemperature
•
Overpressure
•
Loss of ARINC 429 BUS data
•
Any FTIS sensor defective, disconnected, or out of range
INHIBIT MODE
•
Valve failure
In the inhibit mode, electrical power is removed from the system. System operation is stopped, but returns to the normal operating mode after the inhibit conditions cease.
The FTIS is reset through the onboard maintenance system (OMS). The fault is stored in the ICU non-volatile memory (NVM). The FTIS reactivates after a successful built-in test (BIT) and power-up cycle.
The FTIS enters the inhibit mode upon receiving a request from aircraft systems. All FTIS components de-energize in the inhibit mode, but return to an operational state when a change occurs in the status of another system. The following conditions cause an inhibit:
MAINTENANCE MODE
•
Fuel quantity computer (FQC) in refuel mode
•
FQC in defuel mode
•
Main engine start
•
Inflight air-assisted engine start
•
One engine inoperative inflight only
•
Single environmental control system bleed with wing anti-ice on
•
Steep approach with wing anti-ice on
•
Loss of bleed air
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The maintenance mode is accessed through the OMS. To aid in troubleshooting, the FTIS faults occurring during normal operation are recorded in NVM within the ICU. These NVM faults include the values of related parameters at the time of the fault, and are downloaded through the OMS while the aircraft is on the ground. The initiated BIT provides the capability to test the electrical line replaceable units (LRUs).
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-28 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Refueling Defueling
EICAS
Main Engine Start ICU
OMS
Inflight Engine Start One Engine Inoperative Inflight
HMU
18,000 ft
Inhibit
Single ECS Bleed with WAI On Steep Approach With WAI On Loss of Bleed Air
Health Monitoring Mode
Inhibit Mode
Overtemperature Overpressure Loss of A429 BUS
Latched Shutdown
OMS Reset
Start Up
OMS
Sensor Fault
Latched Shutdown Mode
Maintenance Mode
CS1_CS3_4740_007
Valve Fault
Figure 14: Fuel Tank Inerting System Modes Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-29 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
MONITORING AND TEST The following page provides the crew alerting system (CAS) and INFO messages for the control and monitoring system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-30 CS130_47_L2_FTISCI_MonitoringTest.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
CAS MESSAGES
Table 2: INFO Messages Table 1: ADVISORY Message
MESSAGE
MESSAGE
LOGIC
FUEL INERTING FAULT
The fuel tank inerting system fault exists. See INFO message.
47 FUEL INERTING FAULT TEMP ISOL VLV INOP
The fuel tank inerting system is prevented from operating (latched shutdown) until successfully completing a power-up BIT and an IBIT is successfully completed because a fault has been detected with the performance of the FTIS temperature isolation valve.
47 FUEL INERTING FAULT DUAL FLOW SOV INOP
The fuel tank inerting system is prevented from operating (latched shutdown) until successfully completing a power-up BIT and an IBIT is successfully completed because a fault has been detected with the performance of the FTIS dual flow shutoff valve.
Table 2: INFO Messages MESSAGE
LOGIC
47 FUEL INERTING FAULT-FUEL The fuel tank inerting system is INERTING SHUTDOWN prevented from operating (latched shutdown).
LOGIC
47 FUEL INERTING FAULT-FUEL The FTIS oxygen sensor detects INERTING DEGRADED abnormally high oxygen percentage in the nitrogen-enriched air stream or oxygen sensor fault. 47 FUEL INERTING FAULT-FUEL A loss of redundancy in the ICU INERTING REDUND LOSS overtemperature analog backup circuit has occurred, but has not resulted in the FTIS becoming inoperative. 47 FUEL INERTING FAULT-FUEL For inability to confirm IIV or TIV position INERTING INOP or loss of FTIS communication or controller fault. 47 FUEL INERTING FAULT INLET ISOL VLV INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The fuel tank inerting system is prevented from operating (latched shutdown) until successfully completing a power-up BIT and an IBIT is successfully completed because a fault has been detected with the performance of the FTIS inlet isolation valve.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-31 CS130_47_L2_FTISCI_MonitoringTest.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
PRACTICAL ASPECTS INLET ISOLATION VALVE LOCKOUT (TBC) The fuel tank inerting system (FTIS) can be deactivated by locking out the IIV. Access to the IIV is through a small panel in the wing-to-body fairing (WTBF). The valve is locked out by removing the deactivation plug from the stowed position and threading it into the deactivation port. The plug is attached to the valve by a lanyard. Additional procedures require the FUEL INERT CTLR circuit breaker to be locked out.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-32 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
OPE
N
Fuel Tank Inerting Access Panel
Deactivation Port
Deactivation Plug
NOTE Lockout FUEL INERT CTLR circuit breaker.
INLET ISOLATION VALVE
CS1_CS3_4730_007
INLET ISOLATION VALVE
Figure 15: Inlet Isolation Valve Lockout Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-33 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
ONBOARD MAINTENANCE SYSTEM The onboard maintenance system initiated built-in test (IBIT) can be carried out with or without bleed air. The IBIT without bleed air checks all electrically actuated LRUs, however, the IIV, TIV, and DFSOV are only checked for the closed position as these valves are pneumatically operated. To carry out a full test of the IIV, TIV, and DFSOV, the IBIT with bleed air is used.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-34 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
MAINT MENU
MAINT MENU
FTIS ICU-IBIT Without Bleed Air
FTIS ICU-IBIT With Bleed Air IBIT WITH BLEED AIR
IBIT WITHOUT BLEED AIR
1/4
1/4
*** WARNING *** PERSONNEL SHOULD STAY CLEAR OF EXHAUST VENT TO AVOID BREATHING UNHEALTHY FUMES PRECONDITIONS
PRECONDITIONS
1- MAKE SURE THAT AIRCRAFT IS ON-GROUND
1- MAKE SURE THAT AIRCRAFT IS ON-GROUND
2- MAKE SURE AIRCRAFT IS IN MAINTENANCE MODE
2- MAKE SURE AIRCRAFT IS IN MAINTENANCE MODE 3- SET BLEED AIR TO GREATER THAN 20 PSIG
TEST INHIBIT/INTERLOCK CONDITIONS
ON-GROUND STATE MAINTENANCE MODE STATE
TEST INHIBIT/INTERLOCK CONDITIONS
= =
ON-GROUND STATE = MAINTENANCE MODE STATE = BLEED AIR PRESSURE (PSIG) = SNNNNNNN
* INVALID OPERATION * CHECK PRECONDITIONS BEFORE PRESSING "CONTINUE"
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Cancel
Continue
Write Maintenance Reports
Cancel
CS1_CS3_4740_009
PRESSING "CONTINUE" AGAIN WITH INHIBITS CONDITION WILL ABORT THE TEST
Figure 16: Onboard Maintenance System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-35 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
47-20 FUEL TANK INERTING SYSTEM DISTRIBUTION GENERAL DESCRIPTION The nitrogen-enriched air (NEA) distribution system connects the onboard inert gas generation system to the fuel tanks. The dual-flow shutoff valve controls the amount of inert air flow to the fuel tanks. During ground, climb, and cruise operations, the system operates in low-flow mode, generating NEA. During this time, the air space above the fuel is used for NEA storage. This minimizes the amount of NEA required to be generated during descent. During descent, the fuel tank inerting system (FTIS) operates in mid- to high-flow mode, based on the descent rate. When the dual-flow shutoff valve (DFSOV) closes, the valve provides protection against fuel entering the FTIS. The distribution line from the DFSOV to the fuel tank has two springloaded flapper check valves that prevent any fuel that may enter the distribution lines from flowing out of the tank to the FTIS generating system.
The distribution network consists of tubing arranged in symmetrical left and right paths from the entry point. Flow through each path is controlled by flow balancing orifices located in the dry bay. Each of the orifices in the wing and center tanks have unique part numbers and are not interchangeable. They are sized to ensure that NEA flow is evenly distributed to each tank under all flow conditions. To minimize the chance of fuel entering the lines through sloshing or gravitational effects, the tubing is located at high points. The NEA discharge outlets are located as far as possible from the fuel vent line openings, to ensure that the tank is adequately flooded before the NEA flows overboard through the vent scoop. Maintenance port plugs are used to inspect for the presence of fuel in the distribution lines. They also provide access points for conducting pressure tests of the distribution lines and check valves.
The line incorporates a flame arrestor, which prevents fire from reaching the fuel tanks. The flame arrestor consists of a metal canister that contains a grid, specifically designed to eliminate passage of any flame. A J-trap is located immediately downstream of the distribution check valve inside the fuel tank. The J-trap provides a collection point for any fuel that enters the distribution system. This configuration prevents fuel from coming in contact with the distribution check valve flapper, minimizing the chance of fuel entering the FTIS lines. Any fuel in the J-trap is purged from the line during system startup.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-36 CS130_47_L2_FTISD_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
SOL SOL
DFSOV UPSTREAM CHECK VALVE Flame Arrestor
DISTRIBUTION CHECK VALVE
J-Trap
Maintenance Ports
Dry Bay
Center Fuel Tank Left Fuel Tank
CS1_CS3_4720_004
Right Fuel Tank
LEGEND Flow Balancing Orifice – Center Tank Flow Balancing Orifice – Wing Tank Maintenance Port Plug
Figure 17: Fuel Tank Inerting System Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-37 CS130_47_L2_FTISD_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
COMPONENT LOCATION The following components are installed in the distribution system: •
Check valves
•
Flame arrestor
•
Flow balancing orifices
•
Maintenance port plugs
CHECK VALVES The distribution check valve and upstream check valves are located at the front spar outside of the fuel tank near RIB 0. FLAME ARRESTOR The flame arrestor is mounted between the distribution and upstream check valves. FLOW BALANCING ORIFICES There are two flow balancing orifices in each dry bay outboard of RIB 6. One orifice is for the center tank distribution line and one is for the wing tank distribution line. MAINTENANCE PORT PLUGS Two maintenance port plugs are located in the distribution line near the center tank front spar.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-38 CS130_47_L2_FTISD_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
RIB 6L
RIB 0
Center Fuel Tank Front Spar
RIB 6L
To Right Wing
RIB 0 Flame Arrestor
Upstream Check Valve
Center Tank Flow Balancing Orifice
Left Wing Tank Flow Balancing Orifice (right wing tank similar)
Distribution Check Valve
Maintenance Port Plug
J-Trap (ref.)
Maintenance Port Plug
CS1_CS3_4720_005
To Left Wing
Figure 18: Fuel Tank Inerting System Distribution Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-39 CS130_47_L2_FTISD_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-1 CS130_47_ILB.fm
ATA 49 - Airborne Auxiliary Power
BD-500-1A10 BD-500-1A11
49 - Airborne Auxiliary Power
Table of Contents 49-00 APU Construction .........................................................49-2 General Description .........................................................49-2 Left Side View ..............................................................49-2 Right Side View............................................................49-4 Accessory Gearbox Section.........................................49-6 49-10 APU Installation ............................................................49-8 General Description ..........................................................49-8 Component Location .......................................................49-8 Component Information .................................................49-10 Access Doors.............................................................49-10 Access Door Struts ....................................................49-12 APU Drains ...............................................................49-14 Detailed Component Information ....................................49-16 Mounts and Struts......................................................49-16 Eductor ......................................................................49-18 Exhaust .....................................................................49-20 APU Drains Schematic .............................................49-22 49-12 APU Air Intake ............................................................49-24 General Description .......................................................49-24 Air Inlet.......................................................................49-24 APU Air Inlet Door......................................................49-24 APU Air Inlet Door Operation.....................................49-26 Component Location .....................................................49-28 Air Inlet Door ..............................................................49-28 Air Inlet Door Actuator................................................49-28 Air Inlet Door Position Sensor....................................49-28 Controls and Indications .................................................49-30 Detailed Description ......................................................49-32 Monitoring and Tests .....................................................49-34 CAS Messages ..........................................................49-35 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Practical Aspects ........................................................... 49-36 APU Inlet Door Mechanical Rigging.......................... 49-36 OMS APU Inlet Door Rigging.................................... 49-38 APU Door Lockout .................................................... 49-40 Onboard Maintenance System APU Door Inhibit ....................................................... 49-42 49-90 APU Oil System ......................................................... 49-44 General Description ...................................................... 49-44 Component Location .................................................... 49-46 Lubricating Module.................................................... 49-46 Oil Cooler ................................................................. 49-46 Thermostatic/Pressure-Relief Valve.......................... 49-46 Low Oil Pressure Switch ........................................... 49-46 Oil Level Sensor........................................................ 49-46 Gearbox .................................................................... 49-48 Magnetic Chip Collector ............................................ 49-48 Oil Temperature Sensor............................................ 49-48 Air/Oil Separator........................................................ 49-48 Detailed Component Information ................................... 49-50 Lubricating Module.................................................... 49-50 Controls and Indications ................................................ 49-52 Detailed Description ...................................................... 49-54 Monitoring and Tests .................................................... 49-56 CAS Messages ......................................................... 49-57 Practical Aspects ........................................................... 49-58 Oil Servicing .............................................................. 49-58 49-30 APU Fuel System ...................................................... 49-60 General Description ....................................................... 49-60 Component Location ..................................................... 49-62 Fuel Control Unit ....................................................... 49-62
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-i
CS130_M5_49_APUTOC.fm
49 - Airborne Auxiliary Power
Flow Divider Assembly...............................................49-62 Fuel Manifolds and Fuel Nozzles...............................49-62 Plenum Drain Orifice..................................................49-62 Detailed Component Information ....................................49-64 Fuel Nozzles ..............................................................49-64 Detailed Description ......................................................49-66 Fuel Control Unit ........................................................49-66 Fuel Flow Divider and Fuel Flow Divider Solenoid ........................................................49-68 Controls and Indications .................................................49-70 49-40 APU Start and Ignition Systems .................................49-72 General Description .......................................................49-72 Component Location .....................................................49-74 Starter Motor and Clutch Assembly ...........................49-74 Ignition Unit ................................................................49-74 Igniter Plug and Igniter Lead......................................49-74 Detailed Description ......................................................49-76 APU Start Power........................................................49-76 TRU Start ...................................................................49-76 Battery Start ...............................................................49-76 49-50 APU Air Systems ........................................................49-78 General Description .......................................................49-78 Component Location ......................................................49-80 Surge Control Valve...................................................49-80 P2 Sensor ..................................................................49-80 Total Pressure Sensor ...............................................49-80 Differential Pressure Sensor ......................................49-80 Inlet Temperature Sensor ..........................................49-80 Inlet Guide Vane Actuator..........................................49-80 Bleed Air Valve ..........................................................49-80 Detailed Description .......................................................49-82 APU Air Interface .......................................................49-82 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
APU Air Control......................................................... 49-84 Monitoring and Test ....................................................... 49-86 CAS Messages ......................................................... 49-87 49-60 APU Control .............................................................. 49-88 General Description ...................................................... 49-88 APU ECU Control Functions ..................................... 49-90 APU Start .................................................................. 49-92 APU Shutdown.......................................................... 49-94 Component Location ..................................................... 49-96 Exhaust Gas Temperature Thermocouple ................ 49-96 Speed Sensor ........................................................... 49-96 Data Memory Module................................................ 49-96 Electronic Control Unit .............................................. 49-98 Detailed Component Information ................................. 49-100 Data Memory Module.............................................. 49-100 Detailed Description .................................................... 49-102 ECU Power ............................................................. 49-102 ECU Interface.......................................................... 49-104 APU Protective Shutdowns ..................................... 49-106 Monitoring and Tests .................................................. 49-108 CAS Messages ....................................................... 49-109 Practical Aspects ......................................................... 49-110 Onboard Maintenance System................................ 49-110 49-00 APU Operation......................................................... 49-112 General Description ..................................................... 49-112 Controls and Indications .............................................. 49-114 Flight Deck Controls................................................ 49-114 External Controls..................................................... 49-116 Synoptic Page ......................................................... 49-118 Operation .................................................................... 49-120 APU Start ................................................................ 49-120 Normal APU Shutdown ........................................... 49-122
TECHNICAL TRAINING MANUAL
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49-ii
CS130_M5_49_APUTOC.fm
49 - Airborne Auxiliary Power
Emergency APU Shutdown .....................................49-124 Monitoring and Test .....................................................49-126 CAS Messages ........................................................49-127
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-iii
CS130_M5_49_APUTOC.fm
49 - Airborne Auxiliary Power
List of Figures Figure 1: APU Left Side View ...............................................49-3 Figure 2: APU Right Side View.............................................49-5 Figure 3: APU Gearbox ........................................................49-7 Figure 4: APU Enclosure and Installation.............................49-9 Figure 5: APU Access Doors..............................................49-11 Figure 6: APU Access Door Struts .....................................49-13 Figure 7: APU Drains..........................................................49-15 Figure 8: APU Mounts and Struts.......................................49-17 Figure 9: APU Eductor........................................................49-19 Figure 10: APU Exhaust .......................................................49-21 Figure 11: APU Drains Schematic........................................49-23 Figure 12: APU Air Intake.....................................................49-25 Figure 13: APU Air Inlet Door Operation ..............................49-27 Figure 14: APU Air Intake Components ...............................49-29 Figure 15: APU Door Indication............................................49-31 Figure 16: APU Inlet Door Operation....................................49-33 Figure 17: APU Inlet Door Mechanical Rigging ....................49-37 Figure 18: Onboard Maintenance System Inlet Door Rigging................................................49-39 Figure 19: APU Door Lockout...............................................49-41 Figure 20: Onboard Maintenance System APU Door Inhibit..................................................49-43 Figure 21: APU Oil System...................................................49-45 Figure 22: APU Oil Components ..........................................49-47 Figure 23: APU Gearbox Oil Components ...........................49-49 Figure 24: APU Lubricating Module......................................49-51 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 25: APU Status Page Oil Indications ........................ 49-53 Figure 26: APU Oil System Schematic ................................ 49-55 Figure 27: APU Oil Servicing ............................................... 49-59 Figure 28: APU Fuel System ............................................... 49-61 Figure 29: APU Fuel Control Unit, Fuel Manifold, Flow Divider, Fuel Nozzles, and Plenum Drain Valve ............................................ 49-63 Figure 30: Fuel Nozzles ....................................................... 49-65 Figure 31: APU Fuel Control Unit Schematic....................... 49-67 Figure 32: APU Fuel Flow Control ....................................... 49-69 Figure 33: Fuel Low Flow Indication .................................... 49-71 Figure 34: APU Starting and Ignition ................................... 49-73 Figure 35: APU Starting and Ignition Starter Motor, Ignition Unit, Igniter Plug, and Igniter Lead......... 49-75 Figure 36: APU Start Power................................................. 49-77 Figure 37: APU Air System .................................................. 49-79 Figure 38: Air System Surge Control Valve, Bleed Air Valve, Inlet Guide Vane Actuator, and Flow Sensors ............................................... 49-81 Figure 39: APU Air Interface ................................................ 49-83 Figure 40: APU Control........................................................ 49-89 Figure 41: APU ECU Control Functions............................... 49-91 Figure 42: APU Start Sequence........................................... 49-93 Figure 43: APU Shutdown Sequence .................................. 49-95 Figure 44: Data Memory Module, Speed Sensor, and EGT Thermocouples.................................... 49-97 Figure 45: APU Electronic Control Unit Location ................. 49-99
TECHNICAL TRAINING MANUAL
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49-iv
CS130_M5_49_APULOF.fm
49 - Airborne Auxiliary Power
Figure 46: Data Memory Module ........................................49-101 Figure 47: ECU Power........................................................49-103 Figure 48: Electronic Control Unit Interface........................49-105 Figure 49: OMS Fuel Shutoff Valve Test............................49-111 Figure 50: APU Operation ..................................................49-113 Figure 51: APU Flight Deck Controls..................................49-115 Figure 52: APU External Controls ......................................49-117 Figure 53: APU Status Performance Indications ................49-119 Figure 54: APU Start ..........................................................49-121 Figure 55: APU Normal Shutdown .....................................49-123 Figure 56: APU Emergency Shutdown...............................49-125
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-v
CS130_M5_49_APULOF.fm
AIRBORNE AUXILIARY POWER - CHAPTER BREAKDOWN AIRBORNE AUXILIARY POWER CHAPTER BREAKDOWN
Starting and Ignition
Construction
1
5
Air Systems
Installation
2
6
Control
Oil
3
7
Operation
Fuel
4
8
49 - Airborne Auxiliary Power 49-00 APU Construction
49-00 APU CONSTRUCTION GENERAL DESCRIPTION The auxiliary power unit (APU) design consists of two modules: the power section and load compressor, and the gearbox with accessories. LEFT SIDE VIEW The following major components are installed on the APU left side: •
Fuel control unit (FCU)
•
Lube module
•
APU drains
•
Ignition unit
•
Thermostatic pressure-relief valve
•
Air/oil cooler (AOC)
The wiring harness provides an interface between the APU and the electronic control unit (ECU). The starter motor and APU generator have their own wiring harnesses to interface with the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-2
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
Load Compressor
Thermostatic Pressure-Relief Valve
Power Section
Air/Oil Cooler
Gearbox
APU Wiring Harness Connectors
Fuel Control Unit Lube Module
CS1_CS3_4900_004
Ignition Unit
Drains
Figure 1: APU Left Side View Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-3
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
RIGHT SIDE VIEW The following major components are installed on the APU right side: •
APU generator
•
Starter motor
•
Bleed air valve (BAV)
•
Surge control valve
•
Fuel manifold and nozzles
Compartment cooling air is drawn through an eductor that surrounds the exhaust to cool the APU oil and provide compartment ventilation. The bleed air duct provides the interface with the aircraft pneumatic system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-4
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
Fuel Nozzles Eductor
Starter Motor
Exhaust APU Generator
Bleed Air Connection
Surge Control Valve
Bleed Air Valve
CS1_CS3_4900_005 4900 005
Fuel Manifolds
Figure 2: APU Right Side View Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-5
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
ACCESSORY GEARBOX SECTION The accessory gearbox is driven through a quill shaft by the high speed torque of the power section. The gearbox contains a series of spur gears to drive the APU accessories. The following accessories are installed on the gearbox: •
Generator
•
Lubricating module
•
Fuel control unit (FCU)
•
Starter motor
•
Air/oil separator
•
Oil fill port
•
High oil temperature sensor
The gearbox also serves as an oil reservoir for the wet-sump lubrication system. An oil drain plug located near the bottom of the gearbox incorporates a magnetic chip collector that can be removed separately.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-6
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
Air/Oil Separator Starter Motor Lubricating Module
Generator Fuel Control Unit
High Oil Temperature Sensor Magnetic Chip Collector
CS1_CS3_4900_006
Oil Fill Port
Figure 3: APU Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-7
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
49-10 APU INSTALLATION GENERAL DESCRIPTION The auxiliary power unit (APU) sits in an enclosure that consists of five titanium firewalls and a pair of hinged composite access doors. The APU is suspended in the enclosure at a 13° nosedown attitude. The enclosure isolates the APU from the rest of the structure and acts as a load path to support the APU. Seals are provided around the firewall and door interface. There are three access panels on both the right side firewall and left side firewall and two access panels on the top firewall, which are used to gain access to the system components and structures. There is also one access panel located on the aft firewall, to facilitate the removal of the muffler.
COMPONENT LOCATION The APU installation consists of the following components: •
Access doors
•
APU drains
•
Engine mounts and struts
•
Eductor
•
Exhaust
APU compartment cooling and ventilation is provided by the APU air inlet duct when the APU inlet door opens during APU operation. A pair of hinged access doors provides access for servicing and maintenance of the APU. A drain allows accumulated fluids to exit through a small opening in the left hand access door. The electrical wiring harnesses provide the interfaces between the electronic control unit (ECU) and the APU. The APU wiring harness extends from the firewall connector to the front of the APU, and connect all electrical components. APU harness connectors are threaded, stainless steel, and self-locking. A bleed air duct connects the APU pneumatic output to the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-8
CS130_49_L2_APUInstall_General.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Compartment Cooling Air Inlet
Engine Mount APU Drains
Bleed Air Duct
Drain Line Access Panels (8x)
Exhaust
Eductor Struts
CS1_CS3_4911_003
Hold Open Strut (4x)
Figure 4: APU Enclosure and Installation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-9
CS130_49_L2_APUInstall_General.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
COMPONENT INFORMATION ACCESS DOORS The APU doors are attached to the aircraft structure by goose neck hinges. Fireproof door seals located at the interface of the APU firewall and access doors help to prevent fire propagation and provide containment of Halon when extinguisher is activated. The doors are secured closed by two transverse hook latches and four pin latches. A door seal runs the length of the door interface. There are drain holes along the forward edge of the doors. Accumulated fluid from the APU drains exit through a hole near the forward transverse latch.
CAUTION Do not keep the doors in the full open position if winds are more than 65 kt (120 kph). Damage to door structures, hinges, and supports can occur.
NOTE Open the left door first if the two doors are to be opened. Close the right door first when the two doors are open.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-10
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Pin Latch Closed
Pin Latch Open
Drain Hole
Latch Keeper
Hook Latch
Pin Latch Closed
Drain Hole
Pin Latch Open
CS1_CS3_4913_001
Hook Latch
Figure 5: APU Access Doors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-11
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
ACCESS DOOR STRUTS Four struts hold the access doors open. The struts are fixed to the doors and enclosure. When the doors are opened or closed, the struts fold and stow in clips mounted on the doors. A sleeve locks the strut in the extended position when the doors are opened. Each strut has an adjustable rod end to ensure the door is flush with the fuselage.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-12
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
CS1_CS3_4913_002
Sleeve (4x) Strut
(4x)
Figure 6: APU Access Door Struts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-13
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
APU DRAINS The overboard drains are used to drain the APU inlet plenum and exhaust. The drains aid in detecting fuel leaks at the fuel control unit, inlet guide vane actuator and surge control valve, and oil leaks at the load compressor seal. The last drain allows fluid accumulation in the muffler to drain as well as fuel from an unsuccessful start to drain from the combustor plenum drain orifice. The APU Inlet plenum drain tube disposes of the fluids that enter the APU inlet. The fluids drain on the interior of the APU compartment door near a drain hole. All of the APU drains, except for the inlet plenum drain, are grouped at a drain manifold that is supported off the APU structure. The drain manifold provides sealing via a silicone-rubber bathtub seal to the compartment access door. From this manifold, the system drains into a cavity created by the seal and the APU compartment access door. Fluids drain overboard through a drain hole in the compartment access door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-14
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Surge Control Valve and Inlet Guide Vane Actuator
Load Compressor Seal Witness Drain
Combustor Plenum Drain Orifice/ APU Exhaust
CS1_CS3_4912_001
Fuel Control Unit
Inlet Plenum
Figure 7: APU Drains Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-15
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
DETAILED COMPONENT INFORMATION MOUNTS AND STRUTS The APU is suspended in the enclosure at a 13° nosedown attitude by a series of struts and isolators. The system includes four groups of struts that connect to the three mounting points on the APU. The four groups of struts have a suspension system with seven active struts and one catcher strut. A vibration isolator connects each group of struts to the mounting points. The mount system consists of a right side forward isolator with two struts, a left side forward catcher with an adjustable strut, a right side aft isolator with three fixed struts, and a left side aft isolator with one fixed strut, and one adjustable strut. The mount system provides a redundant load path to maintain the APU in position, and provides redundant vibration and shock isolation from the APU to the aircraft. The mount system maintains the APU in position even if one strut fails, or an entire mount is damaged by fire. All isolators react to vertical and lateral loads. The aft isolators also react to thrust loads. The forward left link only reacts to loads in failed condition.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-16
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Strut (7x) Engine Mounts and Isolators (3x)
CS1 CS3 4911 002 CS1_CS3_4911_002
Catcher Strut
Figure 8: APU Mounts and Struts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-17
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
EDUCTOR The eductor provides APU compartment cooling and oil cooling. The eductor uses APU exhaust gas to passively create APU compartment cooling flow. The kinetic energy of the high-velocity exhaust gas draws the flow of lower velocity compartment cooling airstream in a mixing duct. Airflow from the surge system is also introduced in the eductor. Airflow from the compartment is pulled across the oil cooler prior to entering the exhaust stream. The mixed exhaust flow results in lower exhaust system surface temperatures.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-18
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Compartment Cooling Air
Air/Oil Cooler
Eductor
Surge Bleed
CS1_CS3_4921_002
Exhaust
Figure 9: APU Eductor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-19
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
EXHAUST The APU exhaust system is located in the tailcone of the aircraft, aft of the APU compartment. The exhaust has a muffler for noise reduction. The APU exhaust duct attaches to the engine using flexible bellows. The bellows take up the relative motion between the engine and exhaust. A bellows liner provides a smooth surface for the gases to flow through. The bellows bolts to the muffler at the aft firewall. The muffler consists of an outer shell and an inner liner separated by baffles. The inner liner can be removed separately from the outer shell. The muffler is covered by a thermal blanket. The muffler has two lifting points on top for removal. The front muffler face has a fire seal and mounting tabs to mount the exhaust duct to the collar frame at the aft firewall. The muffler has a forward facing drain that expels any fluid accumulation. The drain connects to a flexible tube that drains overboard through the APU access doors The tailpipe is attached to the muffler. The tailpipe is covered by an upper and lower thermal blanket, where it extends through a cutout in the tailcone.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-20
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Upper Thermal Blanket
Tailpipe
Lifting Points
Lower Thermal Blanket
Attachment Point
Liner Muffler
Drain
CS1_CS3_4921_001
Bellows
Figure 10: APU Exhaust Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-21
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
APU DRAINS SCHEMATIC The forward drain disposes of the fluids that enter the APU inlet. The second drain detects degraded seal performance in the fuel control unit (FCU), inlet guide vane (IGVA) actuator, and surge control valve (SCV) actuator. The third drain detects degraded load compressor main shaft seal performance. The aft drain provides a means for disposing of excess fuel in the combustor after an aborted start. It also serves as a drain for any fluid accumulation in the APU exhaust. During normal operations, no fuel is discharged from the aft drain. Telltale drains help find fuel seal failures. They are located in the drain lines from the FCU, IGVA, and SCV. A cap at each drain can be removed and checked for the presence of fuel as part of the troubleshooting procedure for isolating a fuel leak.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-22
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Inlet Guide Vane Actuator
FUEL CONTROL
SURGE CONTROL VALVE Inlet Plenum
Inlet Guide Vane Actuator Telltale Witness Drain
Combustor Case/Educator
Fuel Control Unit
Fuel Control Telltale Witness Drain
Drain Mast Surge Control Valve and Inlet Guide Vane Actuator
Load Compressor Seal Witness Drain
CS1_CS3_4912_002
Surge Control Valve Telltale Witness Drain
Figure 11: APU Drains Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-23
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
49-12 APU AIR INTAKE GENERAL DESCRIPTION AIR INLET The auxiliary power unit (APU) air inlet supplies air for APU operation, cooling, and ventilation. The APU inlet door directs air into the air inlet duct. Cooling air is ducted to the APU compartment by splitting the APU inlet air at the inlet door on the aircraft skin. The inlet duct provides an airflow path to the APU plenum. The compartment cooling air ducting provides cooling air to the APU enclosure. The air flows into the APU compartment, circulates around the APU to vent toxic and combustible gases, before exiting through the exhaust-mounted oil cooler and eductor. During ground operation, the velocity of the hot gas exhaust discharge of the APU through the eductor creates a low-pressure around the APU, pulling the cooling air through the system. During flight operation, ram air at the inlet door provides additional force to move the cooling air. APU AIR INLET DOOR The APU inlet door is an outward opening door that scoops air within the boundary layer and directs the air flow to the APU compressor and oil cooler inlet. In the closed position, the door prevents the APU from windmilling in flight.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-24
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU Air Inlet Door
Compartment Cooling Air Duct
CS1_CS3_4912_003
APU Air Inlet Duct
Figure 12: APU Air Intake Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-25
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU AIR INLET DOOR OPERATION The door opens to 45° for ground operations and 30° in flight. When the door is open, an air gap between the inlet door and the fuselage reduce inlet losses, improve flow stability, and reduce door buffeting. The inlet door actuator moves the lever arm to open and close the inlet door. The inlet door has a mechanical stop that limits the door travel to 50°. The APU electronic control unit (ECU) controls the door position. The actuator motor is powered from the 28 VDC EMER BUS. An air inlet door position sensor provides door positioning information to the ECU for door control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-26
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
45° 30°
DC EMER BUS APU Inlet Door Position Sensor
APU OFF
RUN
START
APU ELECTRONIC CONTROL UNIT
APU Door Actuator
LL PU N R TU
LEGEND APU PANEL
ARINC 429 WOW
CS1_CS3_4912_009
Air/Gnd
Figure 13: APU Air Inlet Door Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-27
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
COMPONENT LOCATION The following components are part of the APU air intake: •
Air inlet door
•
Air inlet door actuator
•
Air inlet door position sensor
AIR INLET DOOR The air inlet door is located on the right side of the vertical stabilizer above the APU compartment. AIR INLET DOOR ACTUATOR The air inlet door actuator is installed on the left side of the air inlet door. AIR INLET DOOR POSITION SENSOR The air inlet door position sensor is installed on the right side of the APU air inlet door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-28
CS130_49_L2_APUAI_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU Air Inlet Door
Air Inlet Door Position Sensor
CS1_CS3_4915_005
Air Inlet Door Actuator
Figure 14: APU Air Intake Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-29
CS130_49_L2_APUAI_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
CONTROLS AND INDICATIONS The APU inlet door status is shown on the STATUS page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-30
CS130_49_L2_APUAI_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
15 °C 15 °C
TAT SAT
ENGINE
103 113 21.8
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L)
APU DOOR Symbol
APU RPM EGT DOOR
100 % 312 °C OPEN
OIL TEMP OIL PRESS OIL QTY
75 °C NORM NORM
OPEN CLOSED OPEN
Condition APU door open. APU door closed. APU door failed open.
TIRE PRESSURE (PSI)
150 224
CLOSED
150
224
224
224
01
01
èè
APU door failed closed. APU door position invalid.
00
00
TEMP
CS1_CS3_4915_006
BRAKE
SYNOPTIC PAGE - STATUS
Figure 15: APU Door Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-31
CS130_49_L2_APUAI_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
DETAILED DESCRIPTION When the ECU receives the APU switch START command, the APU inlet door opens. The ECU supplies power to energize the APU relay and excitation for the air inlet door position sensor. The 28 VDC power is removed from the actuator when the air inlet door position sensor reports the door is in the correct position, based on weight-on-wheels (WOW) sensing. The actuator is driven by the electronic control unit (ECU) logic to the correct position for starting and operating conditions. There are three inlet positions: closed, 30° open for flight, and 45° open for ground operation. If the ECU does not determine that the inlet door is open within 30 seconds after issuing the command, APU starting is inhibited. The ECU supplies 28 VDC to close the inlet door during the normal APU shutdown sequence as well as protective shutdowns. The 28 VDC power is removed when the ECU detects that the inlet door is in the closed position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_49_L3_APUAI_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
EPC 1 APU Relay
INLET DOOR ACTUATOR APU Inlet Door Pwr DC EMER BUS
10
APU ECU Sec DC EMER BUS
10
28 VDC APU ECU
Ground
28 VDC Ground
Open
Close
DC ESS BUS 2
APU OFF
RUN
START
DMC 1 and DMC 2
Start Command Excitation
PULL TURN
INLET DOOR POSITION SENSOR
Feedback
LEGEND
Air/Gnd WOW
EPC
ARINC 429 External Power Center
CS1_CS3_4912_005
APU PANEL
Figure 16: APU Inlet Door Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-33
CS130_49_L3_APUAI_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
MONITORING AND TESTS The following page displays the crew alerting system (CAS) and INFO messages for the APU air intake system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-34
CS130_49_L2_APUAI_MonitoringTests.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
CAS MESSAGES Table 1: CAUTION Message MESSAGE APU DOOR OPEN
LOGIC APU door failed to close when required.
Table 2: INFO Message MESSAGE
LOGIC
49 APU FAULT - APU DOOR APU door failed to open or APU door closed INOP after open or WOW and APU door RVDT fault.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49 - Airborne Auxiliary Power 49-12 APU Air Intake
PRACTICAL ASPECTS CAUTION
APU INLET DOOR MECHANICAL RIGGING The door rigging procedure ensures the door is in the proper closed position in respect to the fuselage and door seal, as well as providing the correct door open positions for ground and flight. Always refer to the Aircraft Maintenance Publication (AMP) for the latest information.
Use caution when manually driving the actuator. A torque of more than 5 in-lb indicates that the actuator has hit its internal stops. Torque in excess of 10 in-lb may damage the APU inlet door actuator, APU inlet door seal or APU inlet door mechanism. Never exceed this torque.
Access to the APU inlet door actuator is through the forward upper access panel on the APU enclosure. The bleed air duct between the bleed air valve (BAV) and the forward firewall must be removed first for access. The door can be manually driven using the 1/4 in. hex manual drive on the door actuator. Turning the manual drive in a clockwise direction to retracts the actuator and opens the door. Turning the manual drive counterclockwise extends the actuator and closes the door. To mechanically rig the door: •
Disconnect the actuator from the inlet door lever arm
•
Manually move the inlet door from fully open to fully closed and ensure the inlet door moves freely
•
Reconnect the lever arm to the inlet door and manually drive the actuator counterclockwise to close the inlet door until it contacts the door seal
•
Manually drive the actuator clockwise to check that the inlet door opens without interference from any components. Manually drive the actuator to close the inlet door.
•
Measure between the inlet door and aircraft skin to make sure a gap of approximately 1 cm (0.4 in) exists.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
Aircraft Skin Inlet Door
2
Door Seal 1
Rod End
1
A torque of more than 5 in-lb indicates that the actuator has hit its internal stops. Torque in excess of 10 in-lb may damage the APU inlet door actuator, APU inlet door seal or APU inlet door mechanism. Never exceed this torque.
2
Clearance between inlet door and fuselage should be 1 cm (0.4 In.)
CS1_CS3_4912_006
NOTE
Manual Drive
Figure 17: APU Inlet Door Mechanical Rigging Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-37
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
OMS APU INLET DOOR RIGGING The door rigging test can be accessed through the onboard maintenance system (OMS) LRU/System operations APU ECU - DOOR RIGGING page. The electrical door rigging test is accomplished after successful completion of the mechanical door rigging. The test lists preconditions that must be followed. The test can only be carried out with the aircraft on the ground and the APU off. When the test is carried out, the door moves. Ensure that personnel and equipment are clear of the door before starting the test. Follow the test instructions on the screen. The door position can be monitored on the OMS and the STATUS synoptic page, and should indicate 45° when open, and 0° when closed. Unsatisfactory results require further mechanical rigging.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
MAINT MENU
RETURN TO LRU/SYS OPS
APU ECU-APU INLET DOOR RIGGING APU INLET DOOR RIGGING
1/10 PRECONDITIONS 1 - SET `APU RUN SW' TO OFF (APU SHUTDOWN) 2 - VERIFY THAT WEIGHT ON WHEELS IS TRUE 3 - DISCONNECT CIRCUIT BREAKER `APU INLET DOOR PWR' 4 - VERIFY `APU DOOR' AND `APU ECU SEC' CIRCUIT 5 - IF THE DOOR IS INHIBITED `OPEN' OR `CLOSED' REFER TO AMP XX-XX-XX TO DEACTIVATE IT RIGGING INHIBIT/INTERLOCK CONDITIONS APU SHUTDOWN: WEIGHT ON WHEELS: INHIBITS OPEN STATUS: INHIBITS CLOSED STATUS:
YYYYY YYYYY L277 B23 L277 B24
PRESS `CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_4912_007
PRESS `CONTINUE' BUTTON TO GO TO NEXT PAGE
Figure 18: Onboard Maintenance System Inlet Door Rigging Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-39
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU DOOR LOCKOUT The APU inlet door lockout pin supports the dispatch of the aircraft with an inoperative APU inlet door actuator or position sensor. The door can be manually driven to the open or closed position from the actuator. The lockout pin is inserted into the actuator arm to lock the door at 0º to keep the door closed if the APU is not going to be operated, or at 30º if the APU is required for flight. The APU can be dispatched for flight with the inlet door in either position. If the door is in the open position, the APU must be operating in flight to prevent the APU from windmilling. If the APU is not operated in flight, the airspeed is limited to 250 KIAS. When the door is locked out in the desired position, the onboard maintenance system (OMS) APU door inhibit is applied.
NOTE The APU can operate with full electrical and pneumatic load on the ground with the door open at 30°.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-40
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
$38 /RFNRXW3LQ
Û&ORVHG 'RRU3RVLWLRQ
Û2SHQ 'RRU3RVLWLRQ
CS1_CS3_4915_004
/RFN3LQ+ROH IRU3RVLWLRQ
Figure 19: APU Door Lockout Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-41
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
ONBOARD MAINTENANCE SYSTEM APU DOOR INHIBIT The onboard maintenance system (OMS) LRU/System Operations APU ECU-Inhibit Tests page is used to inhibit the APU inlet door in the open or closed position when the APU inlet door actuator or position sensor is inoperative. When the APU inlet door has been moved to the desired position, select the applicable door inhibit and then select the Continue soft key. Wait for the TEST OK and either the DOOR SUCCESSFULLY INHIBITED OPEN or the DOOR SUCCESSFULLY INHIBITED CLOSED messages to be displayed. Reactivation of the APU Inlet Door To reactivate the APU inlet door, remove the lock pin from the APU inlet door. If the APU inlet door is not closed, manually drive the door to the fully closed position. On the APU ECU-Inhibit Tests page, select APU DOOR INHIBIT RESET. Confirm the lock pin is removed and select the Continue soft key. Stay on the INHIBIT RESET PAGE for a minimum of five minutes. It is possible that a No response from LRU message may appear. After a successful reset, TEST OK and DOOR INHIBITS SUCCESSFULLY REMOVED messages are displayed. Failure to remain on the INHIBIT RESET PAGE for a minimum of five minutes can result in a failed reset sequence. When the reset is complete, do the OMS APU inlet door rigging as per the aircraft maintenance publication (AMP) to confirm the APU inlet door operation.
CAUTION When the APU inlet door is in the APU DOOR INHIBIT OPEN mode in the OMS, make sure that the door is physically open before an APU start. If the APU is started when the APU inlet door is physically closed while in the APU DOOR INHIBIT OPEN mode, damage to the APU can occur. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-42
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
MAINT MENU
RETURN TO LRU/SYS OPS
MAINT MENU
RETURN TO LRU/SYS OPS
APU ECU-DOOR INHIBITS
APU ECU-INHIBIT TESTS
APU DOOR INHIBIT SELECTION
APU DOOR INHIBIT SELECTION
1/5 *** WARNING *** DO NOT LEAVE THE WRENCH ON THE MANUAL DRIVE. UPON STARTUP, IT COULD SPIN OR BE EJECTED AND INJURE
TEST OK DOOR SUCCESSFULLY INHIBITED OPEN
2/5 SELECT FROM THE FOLLOWING ITEMS:
APU Door Inhibit Open *** NOTE *** WHEN USING THE INLET DOOR ACTUATOR MANUAL CRANK, DO NOT APPLY MORE THAN 10 IN-LBS TO THE 1/4 INCH HEX ON THE BACK OF THE ACTUATOR
APU Door Inhibit Closed
PRECONDITIONS TEST OK
1 - SET `APU RUN SW' TO OFF (APU SHUTDOWN)
APU Door Inhibit Reset
2 - VERIFY THAT WEIGHT ON WHEELS IS TRUE
DOOR SUCCESSFULLY INHIBITED CLOSED
3 - DISCONNECT APU DOOR ACTUATOR POWER TEST INHIBIT/INTERLOCK CONDITIONS APU SHUTDOWN: WEIGHT ON WHEELS: INHIBITS OPEN STATUS: INHIBITS CLOSED STATUS:
INHIBITS/INTERLOCK CONDITION
YYYYY YYYYY L277 B23 L277 B24
- INHIBIT STATUS:
PRESS `CONTINUE' BUTTON TO GO TO NEXT PAGE
NONE
PRESS `CONTINUE' BUTTON TO GO TO CLOSE OUT PAGE
TEST OK
PRESS `CANCEL' BUTTON TO GO TO MAIN MENU DOOR SUCCESSFULLY INHIBITED REMOVED
Write Maintenance Reports
Cancel
Continue
Cancel CS1_CS3_4915_003
Continue
Figure 20: Onboard Maintenance System APU Door Inhibit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-43
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49 - Airborne Auxiliary Power 49-90 APU Oil System
49-90 APU OIL SYSTEM GENERAL DESCRIPTION Oil is drawn from the gearbox through a protective screen into the supply pump. After being discharged from the pump, the oil is regulated to maintain a constant pressure throughout auxiliary power unit (APU) operation. The oil then flows to the thermostatic pressure-relief valve (PRV) and on to the oil cooler. If the oil temperature is low or the oil pressure is high, the oil bypasses the cooler and is distributed to the engine and components. The oil is filtered and distributed to the bearings, gears, and generator. Oil is distributed through internal passages and external lines to the various parts of the APU. The oil is monitored for pressure and temperature. If pressure is too low or temperature is too high, the APU shuts down. The generator is scavenged by elements in the lube module. Oil returning from the generator first passes through inline screens. The pumps discharge oil into a generator scavenge filter and empties it into the reservoir. The generator scavenge and lube module filters have bypass switches that cause the APU to shut down if the filter becomes restricted. Oil from the forward bearings and gearbox returns to the reservoir by gravity. The turbine bearing area is scavenged by a single pump element. Oil passes through an inlet screen, through the pump and into the reservoir. Air and oil are separated by an air/oil separator. The reservoir is vented to the exhaust section which prevents overpressurization of the reservoir during APU operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-44
CS130_49_L2_APUOS_General.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
GEARBOX OIL COOLER STARTER
Air/Oil Separator
EXHAUST
GENERATOR Inlet Screen
LUBE MODULE
Scavenge Return to Gearbox
Oil Level
Inlet Screen
THERMOSTATIC PRESSURERELIEF VALVE
Oil Reservoir
Pressure Scavenge Suction Vent Air
CS1_CS3_4991_003
LEGEND Scavenge Pump Return Screen
Figure 21: APU Oil System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-45
CS130_49_L2_APUOS_General.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
COMPONENT LOCATION
OIL LEVEL SENSOR
The oil system components consist of the following:
The oil level sensor is located on the gearbox, aft of the oil servicing door.
•
Lubricating module
•
Oil cooler
•
Thermostatic pressure-relief valve
•
Low oil pressure (LOP) switch
•
Oil level sensor
•
Gearbox
•
Magnetic chip collector
•
High oil temperature sensor
•
Air/oil separator
LUBRICATING MODULE The lube module is installed on the gearbox. OIL COOLER The oil cooler is an air/oil heat exchanger mounted on the left side of the eductor. THERMOSTATIC/PRESSURE-RELIEF VALVE The thermostatic pressure-relief valve is located on the left side of the gearbox. LOW OIL PRESSURE SWITCH The low oil pressure (LOP) switch is located on the front face of the gearbox next to the lubricating module.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-46
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Oil Cooler
Lubricating Module
Thermostatic Pressure-Relief Valve
Low Oil Pressure Switch
CS1_CS3_4991_002
Oil Level Sensor
Figure 22: APU Oil Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-47
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
GEARBOX The gearbox is mounted to the APU load compressor section. The gearbox holds the APU oil supply. MAGNETIC CHIP COLLECTOR The magnetic chip collector is located on the APU gearbox drain plug. OIL TEMPERATURE SENSOR An oil temperature sensor mounted on the gearbox next to the magnetic chip collector. AIR/OIL SEPARATOR The air/oil separator is located right above and to the left of the lubrication module.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-48
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Air/Oil Separator
Drain Plug With Magnetic Chip Collector Gearbox
Oil Temperature Sensor
CS1_CS3_4991_005
Lubricating Module
Figure 23: APU Gearbox Oil Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-49
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
DETAILED COMPONENT INFORMATION LUBRICATING MODULE The lubricating module is a self-contained unit that provides lubrication and scavenge functions to the generator, gearbox, and main shaft bearings.
Each filter incorporates a filter bypass valve and filter bypass switch to indicate when filter contamination occurs. The filter bypass valve opens when the oil temperature is cold or the filter is clogged. The filter bypass switches signal the ECU when a filter clogs.
The lubricating module incorporates a three-element gerotor pressure pump, a three-element gerotor scavenge pump for clearing oil from the generator, and a single element gerotor scavenge pump for clearing oil from the APU turbine bearing cavity. A pressure regulator maintains a constant lube supply pressure to the engine and generator and a relief valve prevents overpressurization of the oil system. The module incorporates a seal plate to provide sealing features at each oil passageway. This eliminates the need for external tubes at the lubricating module. Deprime Solenoid The deprime solenoid valve is located next to the oil filters on the lubricating module. The deprime solenoid valve opens during cold weather starts to allow air into the oil system. The air helps reduce the drag on the APU during cold start conditions. During APU shutdown, the valve opens to de-oil the lubrication pump. Filters There are two filters installed on the lubricating module. The APU oil filter removes contaminants from the oil before it is distributed to the APU oil system. The a generator scavenge filter removes contaminants from the oil coming from the generator. The oil filter elements are throwaway types contained in a housing that is screwed into the lubricating module housing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-50
CS130_49_L3_APUOS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Lubricating Module
Deprime Solenoid
Seal Plate
Generator Scavenge Filter Element APU Oil Filter Element
CS1_CS3_4991_012
Filter Bypass Switches
Figure 24: APU Lubricating Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-51
CS130_49_L3_APUOS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
CONTROLS AND INDICATIONS The APU oil indications are displayed on the STATUS page. The oil temperature, pressure and quantity are monitored by the ECU. The APU oil pressure shows NORM when the APU is not running.
NOTE The APU oil pressure indication shows NORM when the APU is off. The APU continues to show NORM as the APU starts, and only indicates LOW if there is insufficient oil pressure when the APU is operating above 34% rpm.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUOS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
APU OIL TEMPERATURE Symbol
15 °C 15 °C
TAT SAT
32 °C 144 °C
ENGINE
103 113 21.8
è½½
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L)
Condition APU oil temperature in normal range. Oil temperature at or above high oil temperature amber line threshold. Oil temperature invalid.
APU OIL PRESSURE Symbol
APU EGT DOOR
100 % 312 °C OPEN
OIL TEMP OIL PRESS OIL QTY
75 °C NORM NORM
NORM LOW
Oil pressure below low oil press threshold.
è½
Oil pressure invalid.
TIRE PRESSURE (PSI)
150 224
Oil pressure in normal range.
150
224
224
224
APU OIL QUANTITY Symbol
BRAKE
00
00
TEMP
01
FULL
01
Condition Oil quantity full.
LOW
Oil quantity below low threshold.
NORM
Oil quantity in normal range.
èè½½
Oil quantity invalid.
SYNOPTIC PAGE - STATUS
CS1_CS3_4991_004
RPM
Condition
Figure 25: APU Status Page Oil Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUOS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
DETAILED DESCRIPTION Oil is drawn from the gearbox case reservoir through a coarse inlet screen into the lubrication supply pump elements. The deprime solenoid valve improves APU starting after cold soaked conditions. The deprime solenoid valve is energized to open when the APU start command is received, and the oil temperature is less than -6.6°C (20°F). When the deprime solenoid is opened, air is introduced into the lube pump inlet and the lubrication module loses the ability to circulate oil to the APU. This reduces starter motor drag during cold starting. The valve closes at 60% rpm. On a normal or automatic shutdown, the solenoid is energized at 50% to de-oil the oil system. The oil pressure is regulated to 68 psi. A pressure-relief valve (PRV) opens if the system pressure exceeds 240 psi. The output of the pump is fed to the oil cooler. The thermostatic pressure-relief valve allows the flow of oil to either flow through or bypass the oil cooler, depending on oil temperature. The thermostatic pressure-relief valve allows oil to bypass the oil cooler at temperatures less than 60°C (140°F). Oil flows through the cooler when temperatures are greater than 77°C (170°F). The valve also opens to bypass the oil cooler when there is a differential pressure of 50 psi across the oil cooler assembly. If the thermostatic pressure-relief valve (PRV) fails to close, a high oil temperature shutdown may occur. The oil is filtered and then internally distributed within the gearbox to the oil-cooled generator, engine gears, and bearings. An external line delivers oil to the rear turbine bearing. Both the generator scavenge and APU lube oil filters have filter bypass switches to indicate when filter contamination occurs. On the ground, a contaminated filter causes an APU shutdown and an APU SHUTDOWN advisory message on the engine indication and crew alerting system (EICAS). In flight, an APU caution message appears and the APU does not shut down.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
After lubricating the generator, oil is scavenged from the generator by the generator scavenge elements, and filtered prior to returning to the gearbox. The APU oil drains directly to the gearbox. The turbine bearing compartment oil returns via a dedicated APU scavenge pump element and drains into the gearbox. In scavenging the oil from the bearing cavities, some air is drawn with it since the scavenge pump capacity is greater than the oil flow to the turbine cavity. The scavenged air/oil mixture must be separated and the gear case is vented to prevent the buildup of pressure. The air/oil separator removes the air from the oil as it returns to the gearbox reservoir. The air vents overboard through a line going back to the APU exhaust duct. An oil level sensor, located inside the gearbox adjacent to the oil level sight glass is used to check the APU oil level when the APU is not running. A low oil pressure (LOP) switch monitors the oil flow to the engine bearings. The LOP switch provides protection against low oil pressure conditions. The LOP is a normally closed switch that is checked during prestart and self-test ECU modes. During a prestart built-in test (BIT), the LOP switch has failed if the APU speed is less than 7% and the switch is open. If the LOP switch has failed during the prestart BIT, the APU can start and run but has no LOP protection. An oil pressure auto-shutdown occurs if the APU is on speed, and the oil pressure is less than 35±5 psi for 20 seconds. The oil temperature sensor provides indication, high oil temperature protection, and activation of the deprime solenoid. The normal operating oil temperature is approximately 96°C (205°F) on a standard sea level day. A high oil temperature shutdown occurs when the APU is on speed and the oil temperature is greater than 143°C (290°F) for 10 seconds. A magnetic chip collector collects debris as the oil returns to the gearbox.
TECHNICAL TRAINING MANUAL
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CS130_49_L3_APUOS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Filter Filter Bypass Bypass Switch Switch Filter Bypass Valve
Thermostatic PressureRelief Valve
PressureRegulating Relief Valve
LUBE FILTER
START MOTOR
GENERATOR SCAVENGE FILTER
Turbine Bearing Compartment
Low Oil Pressure Switch
Air Oil Separator Clutch
OIL R COOLE
GENERATOR
Turbine Scavenge Element FCU FWD Bearing Compartment Lubricating Pump
Return Screen
Gen. Scavenge Elements (3x) GEARBOX
Oil Level Sensor Oil Temperature Sensor
Gearbox Air
LEGEND Oil Level Sight Glass
Magnetic Chip Collector Inlet Screen
Regulated Pressure Suction Gearbox Air Scavenge Pressure
CS1_CS3_4991_009
Deprime Solenoid Valve LUBRICATION MODULE
Figure 26: APU Oil System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-55
CS130_49_L3_APUOS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the APU oil system.
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49-56
CS130_49_L2_APUOS_MonitoringTests.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
CAS MESSAGES Table 3: Advisory Message MESSAGE APU OIL LO QTY
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC APU oil low quantity detected.
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUOS_MonitoringTests.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
PRACTICAL ASPECTS OIL SERVICING The gearbox oil supply provides 700 hours of APU operation without oil servicing. The gearbox sump capacity is 7.32 L (7.74 qt). The APU oil servicing is accomplished when the APU access doors are open. The oil quantity can be checked using the sight glass on the side of the gearbox. The sight glass is located on the aft side of the gearbox near the oil service door. The oil level should be checked 30 minutes after shutdown.
NOTE If the oil level is in the ADD mark, oil must be added to the APU gearbox no later than 30 hours of APU operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-58
CS130_49_L2_APUOS_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Oil Service Door
FULL L
1 QT
1 QT T
ADD
ADD D
APU OFF
APU U OFF F
Oil Level Sensor
CS1_CS3_4991_010 CS3 4991 010
FULL
SIGHT GAUGE
Figure 27: APU Oil Servicing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-59
CS130_49_L2_APUOS_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
49-30 APU FUEL SYSTEM GENERAL DESCRIPTION The aircraft fuel system supplies the auxiliary power unit (APU) with low-pressure fuel. Fuel enters the fuel control unit (FCU) through the fuel shutoff solenoid valve and is filtered, pressurized, and metered. The FCU supplies metered fuel to the APU for combustion, and high-pressure fuel to operate the surge control valve (SCV) and inlet guide vane (IGV) actuator. The metered fuel flows from the FCU to the fuel flow divider when the fuel shutoff solenoid valve in the FCU is commanded open by the electronic control unit (ECU). The flow divider distributes the fuel between primary flow and secondary fuel manifolds. The primary flow is used for initial starting of the APU. The secondary flow augments the primary flow for acceleration and on-speed operating conditions. The flow divider solenoid is controlled by the ECU. Ten fuel nozzles inject fuel into the combustion chamber for APU operation. When the ECU receives an OFF signal and the APU has completed a cool-down cycle, the fuel shutoff solenoid valve closes and the APU shuts down. On the bottom of the outer combustion case is a plenum drain orifice. It drains the excess fuel from the combustion chamber in case of a no start condition.
NOTE The plenum drain valve is always open, whether the APU is operating or not. An arrow on the valve body points in the direction of fuel flow toward the overboard drain line. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-60
CS130_49_L2_APUFS_General.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER From Aircraft Fuel System
APU FEED SHUTOFF VALVE
FUEL CONTROL UNIT SOL
SURGE CONTROL VALVE
Primary Fuel Flow
Secondary Fuel Flow
INLET GUIDE VANE ACTUATOR
FUEL NOZZLES
ECU
Discrete Fuel Flow Divider SOL
Combustor
Flow Divider Solenoid Plenum Drain Orifice
CS1_CS3_4931_002
LEGEND
Figure 28: APU Fuel System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-61
CS130_49_L2_APUFS_General.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
COMPONENT LOCATION The fuel system consists of the following components: •
Fuel control unit
•
Flow divider assembly
•
Fuel manifolds and fuel nozzles
•
Plenum drain orifice
FUEL CONTROL UNIT The fuel control unit (FCU) is attached to the lubricating module by means of a quick-release v-groove clamp, and is driven by an oil lubricated splined drive shaft. FLOW DIVIDER ASSEMBLY The flow divider is mounted on the lower left side of the turbine plenum. FUEL MANIFOLDS AND FUEL NOZZLES The primary and secondary fuel manifolds surround the combustor plenum. PLENUM DRAIN ORIFICE The plenum drain orifice is located at the lowest point on the plenum.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-62
CS130_49_L2_APUFS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
Fuel Manifolds
Flow Divider
Fuel Control Unit
Fuel Filter
CS1_CS3_4931_003
Fuel Nozzles
Plenum Drain Orifice
Figure 29: APU Fuel Control Unit, Fuel Manifold, Flow Divider, Fuel Nozzles, and Plenum Drain Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-63
CS130_49_L2_APUFS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
DETAILED COMPONENT INFORMATION FUEL NOZZLES There are 10 dual-orifice fuel nozzles equally spaced and mounted on the combustor plenum. The primary and secondary fuel manifolds route fuel to the fuel nozzles. The fuel nozzles inject metered fuel into the combustor. The primary and secondary fuel inlets have screens to prevent contaminants from entering the spray tip. A locating pin positions the air shroud on the nozzle to ensure proper installation in the engine. Fuel for ignition and initial acceleration is supplied by the primary nozzles. The secondary nozzles provide the additional fuel flow as required for operation up to 100% rpm.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-64
CS130_49_L3_APUFS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
Secondary Fuel
Primary Fuel
Screens Locating Pin Compressor Discharge Air In
Cooling Air Coo Shroud Shro
FUEL NOZZLE
CS1_CS3_4931_007
Nozzle Tip
Figure 30: Fuel Nozzles Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-65
CS130_49_L3_APUFS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
DETAILED DESCRIPTION FUEL CONTROL UNIT The fuel control unit is attached to the lubrication module by means of a quick release v-band clamp. The FCU is driven by an oil lubricated spline drive shaft. The APU fuel system is automatically controlled by the electronic control unit (ECU). The aircraft fuel system supplies the APU with low-pressure fuel from the left engine feed line. Fuel enters the FCU through the inlet filter. The filters trap normal fuel pump wear debris. The entire pump discharge flow passes through the filter and feeds a gear pump that supplies the actuator pressure regulator and the metering valve through the high-pressure fuel filter.
A temperature sensor signals the ECU to adjust flow rates for starting the APU in cold conditions and at high altitude. A flow meter resolver provides fuel flow feedback to the ECU. Once rated speed is reached, fuel flow is modulated to meet the demands of varying pneumatic and electrical loads. When the ECU receives an OFF signal and the APU has completed a cooldown cycle, power is removed from the fuel shutoff solenoid and the APU shuts down. Whenever a protective shutdown is initiated, the signal to the FCU torque motor is automatically removed as well.
A high pressure-relief valve protects the FCU. The relief valve opens at 1150 psi and also relieves pump pressure on all FCU shutdowns. The pressure regulating valve provides 250 ± 25 psi hydraulic pressure used to activate the inlet guide vane actuator and the surge control valve. The fuel metering valve meters fuel flow using a torque motor. The metered fuel is a function of ECU current signal. An increase in current increases fuel flow. Whenever a protective shutdown is initiated, the current to the torque-motor is automatically removed. At 7% rpm, the fuel shutoff solenoid allows fuel to flow to the flow divider and flow divider solenoid. When the solenoid is de-energized, APU operation is terminated.
NOTE If the fuel shutoff solenoid fails to open at 7% speed, a protective shutdown occurs. A leaking fuel shutoff solenoid that allows fuel into the APU before 7% speed causes torching.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-66
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER
SOL
Fuel Temperature Sensor
Fuel Metering Valve
Actuator Pressure Regulator
High-Pressure Fuel Filter Inlet Guide Vane Actuator
FUEL NOZZLES
HighPressure Relief Valve
Flowmeter Resolver
Combustor
Spline Drive Shaft SOL
Fuel Shutoff Solenoid
Drain Orifice
LUBE MODULE GEARBOX
HighPressure Gear Pump
FUEL CONTROL
Fuel Supply Differential Pressure Regulator
FUEL INLET FILTER ELEMENT
Actuator Return Fuel (Low Pressure) Metered Fuel Pump Discharge High-Pressure Regulated Pressure Muscle Press Flow Divider SOL
Flow Divider Solenoid
CS1_CS3_4931_005
LEGEND SURGE CONTROL VALVE
Figure 31: APU Fuel Control Unit Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-67
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER AND FUEL FLOW DIVIDER SOLENOID The flow divider and flow divider solenoid distribute the fuel between primary flow and secondary fuel manifolds. The flow divider solenoid also improves start capability in cold weather. Fuel flows to the primary manifold for APU light-off and initial acceleration. The ECU uses P2, T2, and speed signals to control the fuel flow divider solenoid valve. During an APU start on the ground or at low altitude, the normally open flow divider solenoid is energized closed until 30% rpm. The solenoid also energizes closed when the APU is operating above 23,000 ft or T2 less than 13°C (55°F). This allows the primary fuel manifold pressure to increase above 125 psi and provide better atomization for increased combustion performance. Fuel flow to the secondary manifold is controlled by the flow divider check valve that opens at 120 psi (25-40% rpm). Incorrect flow divider sequencing results in an APU protective shutdown. The flow divider assembly is a line replaceable unit.
NOTE If flow divider sequencing is incorrect, no acceleration, no flame, overtemperature and underspeed, auto-shutdowns could occur.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-68
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER APU RPM < 30% APU Operating Above 23000 ft T2 < 13°c (55°f)
APU ELECTRONIC CONTROL UNIT
From Fuel Control Unit SOL
Solenoid Energized Closed
Flow Divider Check Valve (open at 120 psi (25-40% rpm))
Secondary Fuel Manifold
Primary Fuel Manifold
FUEL NOZZLES
LEGEND
SOL
CS1_CS3_4931_009
Discrete Fuel Flow Divider
Flow Divider Solenoid
Figure 32: APU Fuel Flow Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-69
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
CONTROLS AND INDICATIONS The APU actual fuel flow is compared to the ECU commanded fuel flow. If a disagreement occurs, the LO FLOW indication on the FUEL synoptic page appears and the filter symbol turns from green to yellow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-70
CS130_49_L2_APUFS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
10000 KG 2500 KG
20 °C 3000 KG
20 °C 3000 KG 4000 KG LO FLOW
SYNOPTIC PAGE - FUEL
CS1_CS3_4931_006
APU
Figure 33: Fuel Low Flow Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-71
CS130_49_L2_APUFS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
49-40 APU START AND IGNITION SYSTEMS GENERAL DESCRIPTION The auxiliary power unit (APU) starting and ignition system accelerates the APU to the starting speed and ignites the fuel/air mixture in the combustor. The ECU provides a start command to BPCU 2 when the APU panel START switch is moved to START and the APU inlet door has reached the open position. An APU start uses electrical power supplied by TRU 3 via the TRU start contactor (TSC) or BATT 2 via the battery start contactor (BSC). Once a power source is selected, BPCU 2 closes the APU start contactor (ASC) and the APU starter motor drives the APU through a sprag clutch mounted on the gearbox. The APU start request terminates at 50% rpm and power is removed from the starter. The APU continues to accelerate to a self-sustaining speed. The ignition system provides electrical energy to create a spark at the igniter plug for starting the APU. The ECU controls and powers the ignition system. The ignition system is energized once the start request is made and terminates at 60% rpm. The ignition unit delivers 18,000 V to the igniter plug.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-72
CS130_49_L2_APUSIS_General.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
APU Starter
Igniter
BPCU 2
ASC
Ignition Unit
DMC 1 AND DMC 2
APU ECU BSC
TRU 3
BATT 2
DC BUS 2
DC ESS BUS 2
LEGEND ASC BPCU BSC ECU TSC
APU Start Contactor BUS Power Control Unit Battery Start Contactor Electronic Control Unit TRU Start Contactor ARINC 429
CS1_CS3_4941_002
TSC
Figure 34: APU Starting and Ignition Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-73
CS130_49_L2_APUSIS_General.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
COMPONENT LOCATION The starting and ignition system consists of the following components: •
Starter motor
•
Ignition unit
•
Igniter plug and igniter lead
STARTER MOTOR AND CLUTCH ASSEMBLY The starter motor and clutch assembly is installed on the gearbox, above the generator. The DC starter motor has a brush wear indicator pin. The brush wear indicator pin must be seen through the side window of the indicator housing. If any part of the white indicator pin is visible, the brushes are acceptable for continued use. IGNITION UNIT The igniter unit is located on the bottom of the APU, below the oil cooler. IGNITER PLUG AND IGNITER LEAD The igniter plug is installed on the left side of the combustion section. A single air-gap-type igniter provides ignition for the APU. The igniter consists of an center electrode, an insulator, and an outer shell. The outer shell has cooling holes to aid in reducing tip temperatures. An igniter lead connects the igniter to the ignition unit.
CAUTION Voltage produced by the ignition system is lethal. Caution must be observed when working with the system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-74
CS130_49_L2_APUSIS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
Handcrank
Brush Wear Indicator
Igniter Plug
STARTER MOTOR
CS1_CS3_4941_003
Igniter Lead
Ignition Unit
Figure 35: APU Starting and Ignition Starter Motor, Ignition Unit, Igniter Plug, and Igniter Lead Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-75
CS130_49_L2_APUSIS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
The BPCU commands the ASC open and commands CDC 4 to open the TSC when the APU start from TRU 3 is terminated or aborted.
DETAILED DESCRIPTION APU START POWER The APU can be started using transformer rectifier unit (TRU) 3 or BATT 2. A TRU APU start has priority over a battery APU start if there is no TRU failure in the system. The APU starter motor power is determined by the electrical system configuration. The electrical system applies power to the APU starter motor when BPCU 2 receives an APU start request from the APU ECU. BPCU 2 configures the electrical system for start. The ECU monitors the voltage available to the starter. If an open circuit is detected, a maintenance message is generated. The APU start request discrete signal sent by the APU ECU becomes inactive during a normal start when APU speed exceeds 50% or if the ECU detects an aborted start condition. BPCU 2 aborts and inhibits the APU start function if the start request from the ECU is not terminated after 60 seconds. Two failed APU start attempts followed by one successful attempt are permitted during normal operation. Do not perform more than three starts/start attempts in one hour. A two-minute delay must be observed between start attempts. A cooldown period is required for the TRU after three failed starts. The battery must cooldown and recharge after three failed start attempts. If any power contactor failure is detected, or there is an AC power transfer, the APU start is aborted.
BATTERY START An APU start from battery occurs when both batteries are available and the TRU 3 start is inhibited. Two batteries are in use during the APU starting interval. Battery 2 is dedicated to the APU starter, and battery 1 powers the aircraft. An APU start from battery 2 is inhibited when it is the only source of power on the aircraft or any associated APU start power contactors are failed. For a battery start, BPCU 2 commands battery line contactor (BLC) 2 to open. Following confirmation the BLC 2 status is open, BPCU 2 requests CDC 2 to close the BSC. CDC 2 monitors the status of BSC contactor and reports it to BPCU 2. When the BSC is confirmed closed, BPCU2 commands the ASC to close. When the APU start from battery 2 is terminated or aborted, the BPCU commands the ASC open and commands CDC 2 to open the BSC.
TRU START APU starts from TRU 3 is inhibited if any TRU is not powered. For a TRU APU start, BPCU 2 commands CDC 4 via CDC 2 to close the TRU start contactor (TSC). CDC 4 commands the TSC coil to close. CDC 4 monitors and sends the status of the TSC contactor to BPCU 2. Following confirmation the TSC is closed, BPCU 2 closes the APU start contactor (ASC).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-76
CS130_49_L3_APUSIS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
Start Power Selection Logic
EPC 3
TRU 1 TRU 2 TRU 3
TRU
BATT 1 BATT 2
BATT
APU START
TSC TRU 3
AC ESS BUS
TSC Status CDC 4 CDC 4-4-10 TSC Ctrl DC BUS 2
LEGEND ASC BLC BPCU BSC CDC EPC RLC TSC
SSPC 3A RLC
Logic: Advanced
APU Start Contactor Battery Load Contactor BUS Power Control Unit Battery Start Contactor Control and Distribution Cabinet Electrical Power Center Ram Air Turbine Load Contactor TRU Start Contactor ARINC 429 TTP
EPC 2 BSC BATT 2
BATT DIR BUS 2
ASC
APU STARTER MOTOR
BSC Status CDC 2 CDC 2-7-12 BSC Ctrl
APU RUN
START
SSPC 3A
Logic: Advanced PULL TURN
APU ECU
BLC
BPCU 2 ASC Status APU Start Command
APU PANEL
APU Starter Voltage Sensing
CS1_CS3_4941_005
OFF
DC ESS BUS 2
Figure 36: APU Start Power Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-77
CS130_49_L3_APUSIS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
49-50 APU AIR SYSTEMS GENERAL DESCRIPTION The auxiliary power unit (APU) supplies compressed air for air conditioning and main engine starting from the load compressor. The pneumatic output is through the bleed air valve (BAV).
The BAV is a pneumatically actuated butterfly valve used to control bleed flow to the airplane. The BAV is normally in the closed position and uses load compressor bleed air controlled by a solenoid to provide the muscle for valve opening.
The load compressor takes air from the same inlet as the power section compressor. It compresses the air and supplies it to the aircraft through a bleed air valve. The bleed air valve shuts off the air flow from the load compressor when there is no demand placed on the APU by the aircraft pneumatic system. The load compressor provides the maximum amount of air flow that the aircraft demands. Inlet guide vanes (IGVs), located between the load compressor and the air inlet, match the compressor flow to the demand. The inlet guide vanes are positioned by an inlet guide vane actuator (IGVA). The ECU controls the IGVA and surge control valve (SCV), using environmental data from P2 (pressure) and T2 (inlet temperature), load compressor output (Pt and P), and load demand from the aircraft. When there is no demand, the ECU closes the IGVs to the lowest possible setting, so the load compressor imposes a minimum load on the power section. In this condition, the bleed air valve is closed and all air is directed through the SCV. The SCV prevents load compressor surge for all operating modes, altitudes, and inlet temperatures. Surge is prevented by opening the SCV when the aircraft pneumatic system takes little or no airflow, or when the BAV is closed. During the APU shutdown sequence, the ECU initiates pneumatic unloading by closing the IGVs and BAV, and opening the SCV.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-78
CS130_49_L2_APUAS_General.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
T2
Fuel In Out
To ECU P2
IGVA
To ECU
Oil Cooler
Inlet Guide Vanes
ECU
Sensor PT Sensor SOL
To Aircraft Pneumatic System
Surge Bleed Air Exhaust LEGEND
Fuel In Out
SCV
BAV
To ECU
BAV ECU IGVA LVDT PT P2 SCV TM T2 ¨3
Gearbox Vent APU Inlet Air APU Combustion Air Bleed Air Fuel Bleed Air Valve Electronic Control Unit Inlet Guide Vane Actuator Linear Variable Differential Transformer Total Pressure Inlet Pressure Sensor Surge Control Valve Torque Motor Inlet Temperature Sensor Differential Pressure Discrete
CS1_CS3_4951_008
¨P
Air Inlet
Load Compressor Outlet
Figure 37: APU Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-79
CS130_49_L2_APUAS_General.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
COMPONENT LOCATION
INLET GUIDE VANE ACTUATOR
The air system consists of the following components:
The inlet guide vane actuator is located below the load compressor section.
•
Surge control valve
•
P2 sensor
BLEED AIR VALVE
•
Total pressure sensor
•
Differential pressure sensor
The bleed air valve is on the right side of the APU compartment in the bleed air duct.
•
Inlet temperature sensor
•
Inlet guide vane actuator
•
Bleed air valve
SURGE CONTROL VALVE The surge control valve is located in the surge bleed duct on the right side of the APU. P2 SENSOR The P2 sensor is installed on the right side of the inlet plenum. TOTAL PRESSURE SENSOR The total pressure (Pt) sensor is located on the bottom of the APU compressor plenum. DIFFERENTIAL PRESSURE SENSOR The differential pressure (P) sensor is located on the bottom of the APU compressor plenum. INLET TEMPERATURE SENSOR The inlet temperature (T2) sensor is located on the bottom of the APU compressor plenum.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-80
CS130_49_L2_APUAS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
ǻP Sensor Bleed Air Valve
PT Sensor
P2 Sensor Surge Control Valve
LEGEND PT Sensor P2 Sensor T2 Sensor ǻP Sensor
Total Pressure Sensor Inlet Pressure Sensor Inlet Temperature Sensor Differential Pressure Sensor
Inlet Guide Vane Actuator
CS1 CS3 4951 003 CS1_CS3_4951_003
T2 Sensor
Figure 38: Air System Surge Control Valve, Bleed Air Valve, Inlet Guide Vane Actuator, and Flow Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-81
CS130_49_L2_APUAS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
DETAILED DESCRIPTION APU AIR INTERFACE The APU electronic control unit (ECU) communicates with the aircraft systems via the data concentrator module cabinet (DMC) 1 and DMC 2 over ARINC 429 BUSES.
The pneumatic output of the APU is supplied to the aircraft through a bleed air valve. The bleed air valve is a solenoid controlled valve, receiving 28 VDC form the ECU. A microswitch in the valve provides an indication when the valve is open.
The APU generator control unit (GCU) supplies electrical load information to the ECU and receives a ready-to-load signal when the APU is on-speed. The ECU receives environmental control system (ECS) demands from the integrated air system controller (IASC) 1 and IASC 2. The ECS demands are based on air data system information and whether the aircraft is on the ground or in flight. The left electronic engine control (EEC) and right EEC provide main engine start (MES) and MES termination commands. The control of the pneumatic output is based on ECS or MES demands as well as the APU inlet temperature (T2) and inlet pressure (P2). The pneumatic output is based on the position of the inlet guide vanes. The IGVA is positioned by a torque motor driven by the ECU. The IGVA provides position feedback to the ECU. When the pneumatic demand changes, due to ECS pack valve closures or the termination of MES, the surge control valve opens to prevent an APU surge. The surge control valve is positioned by a torque motor driven by the ECU. The ECU receives position feedback from the SCV. The SCV position is based on the load compressor total pressure (Pt) and differential pressure (∆P). The Pt, ∆P, and P2 sensors receive 10 VDC excitation from the ECU and supply outputs proportional to the sensed pressure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-82
CS130_49_L3_APUAS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
AIR COCKPIT
FWD
CABIN
C
AFT
H
FWD
MAN TEMP
CARGO
VENT
OFF
LO HEAT
OFF
HI HEAT
ON
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
FAIL
IGVA
Command Position
SCV
28 VDC Valve Position
BAV
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
Command Position
FAIL OFF R BLEED FAIL
APU BLEED
OFF
FAIL OFF
IASC 1 and IASC 2 APU GCU
LEGEND
DMC 1 and DMC 2
∆P
LEFT EEC and RIGHT EEC
10 VDC Excitation
PT P2
Air Data Information
T2 Air/Gnd
APU ELECTRONIC CONTROL UNIT
WOW
BAV DMC EEC GCU IASC IGVA PT P2 SCV T2 ∆P
Bleed Air Valve Data Concentrator Unit Module Cabinet Electronic Engine Control Generator Control Unit Integrated Air System Controller Inlet Guide Vane Actuator Total Pressure Inlet Pressure Sensor Surge Control Valve Inlet Temperature Sensor Differential Pressure ARINC 429
CS1_CS3_4950_001
AIR PANEL
Figure 39: APU Air Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-83
CS130_49_L3_APUAS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
APU AIR CONTROL The load compressor output is regulated by variable IGVs located just upstream of the load compressor impeller. Output is based on aircraft environmental control system (ECS), main engine start (MES), and electrical load demands as indicated by aircraft signals to the APU electrical control unit (ECU). The IGV position is a function of P2, T2, and load demand. During an APU start, the IGVs are closed to 15° up to 60% rpm and then opened to 22° if altitude is below 25,000 ft. This removes any pneumatic load from the load compressor. Above 25,000 feet, the IGVs are held at 15°.
The surge control valve (SCV) ducts air from the load compressor to the APU exhaust. The SCV makes sure there is a minimum flow of air through the load compressor. The surge margin set point is the minimum quantity of air that should flow through the load compressor to prevent load compressor surge. The ECU calculates the surge margin set point using: •
Inlet pressure (P2)
•
Inlet temperature (T2)
To minimize fuel consumption, when there is no demand, the ECU closes the IGVs to the lowest possible setting that provides a constant airflow through the compressor during all loading conditions so the load compressor imposes minimum load on the power section. The ECU logic for determining IGV position includes a compensation factor for load compressor deterioration, based on the APU operating hours recorded by the data memory module (DMM).
•
IGV position
•
Bleed mode
•
Air/ground mode
To optimize fuel consumption in the ECS mode, the pneumatic load provided by the APU is matched to the aircraft requirements. The pressure and flow requirements vary with ambient conditions and whether the aircraft is on the ground or in flight.
During MES, the APU ECU opens the IGVs to 90° and modulates the SCV to provide maximum bleed for engine starting. The priority is given to the pneumatic load during engine start. If required, some APU generator load sheds.
As the combined loads of the generator and the load compressor are imposed on the power section, the speed governing function of the ECU responds by increasing fuel flow. Under some extremely high combination loads or with significant deterioration of the power section, the turbine temperature limit is reached.
To prevent damage to air conditioning components as well as overpressure and overtemperature conditions, the APU software logic limits the APU bleed pressure and temperature to: •
65 psi maximum bleed pressure
When a load change results in a temperature higher than its limit, the ECU moves the IGVs toward a more closed position, reducing load compressor airflow, and thus reducing the pneumatic load and decreasing the turbine temperature. The electrical load remains unchanged and has priority over the pneumatic load.
•
246ºC (475ºF) maximum bleed temperature
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The ECU uses total pressure (Pt) and differential pressure (P) to maintain the load compressor output at or above the stall margin set point.
In the event of a failure, the ECU opens the SCV and closes the IGVs. This gives priority to electric operation over bleed air demand in the event of a surge system malfunction.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-50 APU Air Systems
Table 4: APU AIR System Operating Modes MODE Startup
READY-TO-LOAD (RTL) (Generator only)
Environmental Control System
Main Engine Start
ELECTRICAL LOAD AVAILABILITY
DEFINING CONDITIONS •
APU ON signal present
•
MOMENTARY initiates start
START
•
Speed >95% +2 seconds
•
BLEED-AIR COMMAND OFF
•
Not available
PNEUMATIC SYSTEM STATUS •
IGVs positioned engine start
during •
Timed acceleration starting logic applies
•
SCV fully open
•
Mode ends when APU reaches 95% speed
•
Bleed air not available
•
•
IGVs positioned at 15° for altitude >25,000 ft
•
SCV fully open
Bleed air not available, only because APU is not selected as bleed air source at conditions where bleed air would normally be available
•
Bleed air available
•
Electrical load has priority
•
IGVs set to ECS position schedule • by logic in the ECU, as a function of inlet temperature (T2), inlet pressure (P2), A/C configuration, and flow mode
•
Surge control system active
signal
Speed >95% +2 seconds
•
BLEED-AIR COMMAND ON
•
Main engine start is not requested
•
One or both packs on
•
Altitude <23,000 ft
Available
Available
at
15°
•
Speed >95% +2 seconds
•
Bleed air available
•
BLEED-AIR COMMAND ON
•
IGVs set at 90° (full open)
•
Either left or right main engine start signal is true
•
Surge control system active
•
Altitude <23,000 ft
Available
COMMENT
•
If power capability is exceeded, pneumatic load is cut back, as needed, by closing the IGVs
Pneumatic power has priority. If power capability is exceeded, load shed is sent to airplane to reduce electrical load, so that IGVs do not have to be cut back
CS1_CS3_TB49_001
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-85
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49 - Airborne Auxiliary Power 49-50 APU Air Systems
MONITORING AND TEST The following page provides the crew alerting system (CAS) messages for the APU air system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-86
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49 - Airborne Auxiliary Power 49-50 APU Air Systems
CAS MESSAGES Table 5: CAUTION Message MESSAGE APU BLEED FAIL
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC Failure of APU providing bleed air when requested or bleed leak detection fail.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
49-60 APU CONTROL GENERAL DESCRIPTION The auxiliary power unit (APU) electronic control unit (ECU) controls and monitors the APU operation and provides an interface with the aircraft. The APU ECU uses inputs from the exhaust gas temperature (EGT) thermocouples and the speed sensor to control the start and acceleration sequence, speed governing, load sequencing, and the shutdown sequence. The ECU has field-loadable software.
The ECU shuts down the APU when the built-in test (BIT) identifies conditions that can damage the engine or the aircraft for continued APU operation. The ECU also has the authority to inhibit starting of the APU when the BIT indicates APU operation may cause damage.
The EGT thermocouple senses the temperature of the exhaust gas that exits out of the turbine section. The speed sensor provides two independent speed signals to the ECU. The APU ECU operates on 28 VDC power provided by DC ESS BUS 2, with the DC EMER BUS supplying a secondary source of power. The APU is started using the APU switch on the APU panel. The ECU controls the ignition and fuel automatically, based on environmental conditions. The APU runs at a variable speed (91-98%) on the ground, and is governed to 100% rpm in flight. The ECU sends a ready-to-load signals to allow electrical or pneumatic loading when the APU is approaching its governed speed. The data memory module (DMM) is an electronic data storage device mounted on the APU and is used for recording parameters, such as APU operating hours, starts, number of unsuccessful starts, and the APU serial number. The ECU monitors the APU components for faults, recording, and reporting. A failed ECU causes an immediate APU shutdown and an APU SHUTDOWN EICAS advisory message to be displayed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-88
CS130_49_L2_APUControl_General.fm
49 - Airborne Auxiliary Power 49-60 APU Control
DC EMER BUS
APU OFF
RUN
EICAS
DC ESS BUS 2 START
SYN ELECTRONIC
PULL TURN
OMS
CONTROL UNIT
AHMS Speed Sensor
APU PANEL
INLET GUIDE VANE ACTUATOR
Oil Cooler
Exhaust Gas Temperature Thermocouples
DMM GENERATOR
G e a r b o x
IGN UNIT FUEL CONTROL UNIT
DMM IGN
BLEED AIR VALVE
SURGE CONTROL VALVE
Data Memory Module Ignition Unit Discrete Fuel RS-485
APU Fuel Feed Line
CS1_CS3_4961_008
LEGEND
To ACFT Pneumatic System
Figure 40: APU Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-89
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49 - Airborne Auxiliary Power 49-60 APU Control
APU ECU CONTROL FUNCTIONS Prestart Sequence
Normal Shutdown
The ECU performs integrity checks to verify that the APU is safe to start and operate. The requirements for an APU start are:
The ECU initiates APU shutdown by injecting a false overspeed signal into the programmable logic device (PLD), which provides overspeed protection. The ECU uses two independently operating PLD devices. The function of each overspeed PLD is to determine if an APU overspeed condition exists and cause a shutdown.
•
ECU BIT self-test passed
•
ECU BIT test of sensors and controls
•
APU door full open signal
Start Sequence When the inlet door reaches the proper position, the ECU sends a start enable command to BPCU 2 to power the APU starter. The ECU powers the ignition unit, and then energizes the fuel solenoid. As the APU accelerates to operating speed, the ECU controls fuel flow via the fuel control unit torque motor as a function of acceleration, fuel schedule, and EGT. The ignition unit and starter command are deactivated when the speed thresholds are attained. On-Speed Governing Exhaust gas temperature and engine speed are continuously monitored by the ECU. Once on-speed, the ECU maintains engine speed at 100% rpm. On the ground, the APU is governed between 91% and 98% rpm based on environmental conditions. Ready-to-load signals are provided for electrical and pneumatic systems.
The primary mode for shutdown is by removing power from the fuel solenoid. Each PLD is capable of shutting down the APU in the event of an overspeed. In order to limit potential fuel solenoid dormancy, at every normal APU shutdown, the APU fuel shut-off solenoid is exercised, alternating the commanding PLD at each APU shutdown. Protective Shutdowns The ECU shuts down the APU when the BIT identifies conditions that can damage the engine or the aircraft. APU protection is provided by logic schemes that determine whether the APU is safe to operate and protective shutdown logic that shuts down an operating APU. Fault Detection The ECU monitors the APU components for faults, recording and reporting. A failed ECU causes an immediate APU shutdown and an APU SHUTDOWN EICAS advisory message to be displayed. The ECU reports faults to the onboard maintenance system (OMS).
Pneumatic Loading Pneumatic loading for the aircraft environmental control system and main engine start operation are provided by the APU through the ECU control of the IGVs. The ECU positions the IGVs in response to bleed demands. The ECU regulates the APU pneumatic output by sensing the APU EGT and comparing it to a predetermined schedule within the ECU. APU surge is prohibited through the ECU control of the SCV actuator. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-90
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49 - Airborne Auxiliary Power 49-60 APU Control
Prestart Sequence Start Sequence On-speed Governing APU ECU
Pneumatic Loading Normal Shutdown Protective Shutdown Fault Detection
CS1_CS3_4961_009
APU ECU
Figure 41: APU ECU Control Functions Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-91
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49 - Airborne Auxiliary Power 49-60 APU Control
APU START The APU switch signals the ECU when it is selected to the START position. The ECU opens the APU feed shutoff valve and the APU inlet door. When the air inlet door is fully open, the inlet door position sensor sends a door open signal to the ECU. The APU start sequence is as follows: •
0% rpm, the ECU energizes the starter and ignition circuit
•
0% rpm, the IGVs positioned to 15°
•
7% rpm, the fuel solenoid valve opens
•
50% rpm, the starter de-energizes
•
60% rpm, the ignition unit de-energizes
•
On ground, ready to load signal is 3% below governed speed signal (88% rpm to 95% rpm)
•
In flight, ready to load signal is 95% rpm
•
91-98% rpm, governed speed range on ground
•
100% rpm, governed speed in flight
•
106% rpm, overspeed shutdown is initiated
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
APU OFF
RUN
START
106%
Maximum Speed
100% 98% 95%
100% Governed Speed (Flight and Groomer Mode) Ready to Load (Flight)
91%
Electrical and Pneumatic Loading on Ground (Ready to Load 88% to 95%)
PULL TURN
60%
Ignition Unit De-energized
50%
Starter De-energized
7% 0%
Fuel Solenoid Valve Opens Ignition Unit Energized
CS1_CS3_4961_015
• APU switch to START position and release to RUN • APU feed SOV open • Air inlet door open • Starter energized
Figure 42: APU Start Sequence Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-93
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49 - Airborne Auxiliary Power 49-60 APU Control
APU SHUTDOWN
Cooldown Cycle
The ECU controls the APU shutdown. There are two types of shutdown conditions: normal and protective.
The cooldown period prevents oil coke at the turbine bearing and fuel nozzles. During cooldown, the ECU:
Normal Shutdown A normal shutdown is initiated from the APU panel when the APU switch is moved to the OFF position. The normal shutdown sequence is as follows:
•
Removes the bleed air ready-to-load signal
•
Closes the bleed air valve
•
Closes the inlet guide vanes to 22°
•
Opens the surge control valve
•
28 VDC ON signal removed from ECU
•
De-energizes the generator
•
ECU receives the OFF signal
•
Starts the 60 second timer
•
Generator ready to load signal is removed
•
60 second cooldown period starts
•
After cooldown, the ECU tests the overspeed circuit closes the FCU shutoff valve and removes power from the metering valve -
Protective Shutdown A protective shutdown is similar to the normal shutdown but has no cooldown cycle.
ECU tests the overspeed circuit which closes the fuel solenoid shutoff valve and shuts down the APU
NOTE
-
During this time, the ECU tries to maintain the APU speed by modulating the FCU torquemeter current
The APU fuel shutoff valve and air inlet door close for both normal and protective shutdowns.
-
If the APU continues to run, the fuel shutoff solenoid valve has failed to close
-
If the fuel solenoid shutoff valve has failed, the ECU reduces the FCU torquemotor current to 0 to shut down the APU
•
20% speed, the APU inlet door is commanded closed
•
7% rpm, the APU fuel shutoff valve (SOV) closes
•
At less than 7% speed, an APU restart can be initiated
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-94
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49 - Airborne Auxiliary Power 49-60 APU Control
APU Operating
60 Second Cooldown On Speed
100%
APU RUN
START
APU Switch Off
LL PU RN U T
• Ready-to-load signal removed • Bleed air valve closed • Inlet guide vanes closed • Surge control valve open • Generator off • 60 second timer starts for cooldown
20% APU Inlet Door Commanded Closed
20%
7% APU Feed Shutoff Valve Closes
7%
APU Restart Possible 0%
CS1_CS3_4961_018
OFF
Figure 43: APU Shutdown Sequence Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-95
CS130_49_L2_APUControl_General.fm
49 - Airborne Auxiliary Power 49-60 APU Control
COMPONENT LOCATION The control system consists of the following components: •
Exhaust gas temperature thermocouple
•
Speed sensor
•
Data memory module
•
Electronic control unit
EXHAUST GAS TEMPERATURE THERMOCOUPLE The APU consists of two exhaust gas temperature thermocouples probes that are located below the APU turbine casing and extend into the exhaust gas stream inside the APU. SPEED SENSOR The speed sensor is located on the bottom of the load compressor case. DATA MEMORY MODULE The data memory module is installed on the left side of the APU air inlet plenum.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
DMM
Speed Sensor
CS1_CS3_4961_006
EGT Thermocouples
Figure 44: Data Memory Module, Speed Sensor, and EGT Thermocouples Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-97
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49 - Airborne Auxiliary Power 49-60 APU Control
ELECTRONIC CONTROL UNIT The electronic control unit (ECU) is located in the mid equipment bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
CS1_CS3_4961_005
MID EQUIPMENT BAY AFT RACK
APU ECU
Figure 45: APU Electronic Control Unit Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-99
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49 - Airborne Auxiliary Power 49-60 APU Control
DETAILED COMPONENT INFORMATION CAUTION
DATA MEMORY MODULE The data memory module (DMM) interfaces with the APU ECU. The APU ECU manages the communication between the OMS and DMM. The DMM stores information regarding the health monitoring and life usage of the APU. The DMM interfaces with the ECU via a serial link and stores the following information: •
APU serial number and model number
•
APU operating hours and starts
•
APU low cycle fatigue and creep life
•
APU accumulated aborted starts
Do not remove the APU electronic control unit (ECU) and data memory module (DMM) at the same time. Always do the APU BIT check between replacement of each of the units or data may be erased.
At power-up, the ECU reads the DMM and performs the following: •
If the information is the same as the ECU, power-up initialization occurs
•
If the information is different, the ECU record is updated from the DMM
•
Upon completion of an APU cooldown cycle, the DMM is automatically updated from the ECU
The APU data is also stored in the ECU in case of DMM failure. Replacement DMMs are delivered with all zeros stored in their memory. Once recognized, the ECU records in memory that a DMM change has occurred. The new DMM is then loaded with all the totaled APU data that is stored in the ECU. If the ECU is replaced, the DMM retains all the APU data and automatically re-initializes the new ECU with the old accumulated APU data.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
APU
OMS
DMC 1 and DMC 2
Data Memory Module
CS1_CS3_4961_004
ELECTRONIC CONTROL UNIT
LEGEND ARINC 429 RS 485 Discrete
Figure 46: Data Memory Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
DETAILED DESCRIPTION ECU POWER The ECU primary power is supplied by DC ESS BUS 2. The ECU also has a secondary power input that is connected to the aircraft DC EMER BUS through the APU relay controlled by the ECU. The dual input power configuration allows the ECU to select the highest voltage power source, remain powered ON whenever the DC EMER BUS is switched on, and allows the DMC to continuously communicate with the ECU whenever the aircraft is powered. APU Start During normal operation, the ECU is powered on via the DC ESS BUS 2 when the aircraft battery switch is closed. After the APU switch is turned to RUN and the ECU completes successful prestart BIT, the ECU applies power to the APU relay and provides secondary power to the ECU via the DC EMER BUS. APU Shutdown On the ground, once the APU switch is turned to OFF and the APU 60 second cooldown, shutdown, and memory management activities are completed, the ECU de-energizes the relay. The APU ECU power is only supplied by the DC ESS BUS 2. In flight, the relay remains activated at APU shutdown to retain power input redundancy.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
EPC 1 APU Relay
APU Inlet Door Pwr DC EMER BUS
10
APU ECU Sec DC EMER BUS
10
APU ECU
28 VDC Ground
APU RUN
START
Start Command
PULL TURN
APU PANEL
CDC 2-8-1 APU ECU PRI DC ESS BUS 2 SSPC 10A Logic: Always On
CS1_CS3_4961_010
OFF
Figure 47: ECU Power Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
ECU INTERFACE The ECU communicates with the IASC to determine the APU bleed mode required. The ECU positions the IGVs and SCV based on this information.
The groomer mode is entered when the maintenance outlets in either galley 1 or 4 are used. In the groomer mode, the APU is governed to a frequency of 400 Hz.
The ECU receives DC electrical power from DC ESS BUS 2 and the DC EMER BUS. The ECU receives a signal from the fire detection and extinguishing (FIDEX) system when an APU fire is detected. When pressed, the APU FIRE pushbutton annunciator (PBA) also closes the shutoff valve (SOV), independent of the ECU. The ECU controls the APU fuel shutoff valve during normal operation. The ECU provides information to the EICAS and synoptic pages and receives input signals from aircraft systems via an ARINC 429 BUS from the data concentrator unit module cabinets (DMCs). The DMCs provide: •
Electrical load
•
Electrical load shed requests
•
Main engine start requests
•
Engine status
•
Weight-on-wheels (WOW) status
•
Air data
•
APU switch commands
Unless otherwise specified, when duplicate data is available from both DMCs, the ECU selects DMC 1 as the primary data source. The ECU controls and receives position feedback from APU inlet door actuator. The APU can be shut down externally by a switch located in the external service panel or from the switch in the APU compartment.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
LEGEND ADSP AGCU APU CDC DMC ECU EPC FADEC IASC LGSCU
Air Data System Probe Auxiliary Generator control Unit Auxiliary Power Unit Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Control Unit Electronic Power Control Full Authority Digital Engine Control Integrated Air System Controller Landing Gear and Steering Control Unit
EPC 1 APU RELAY
APU ECU
Secondary Power Relay Low
CDC 2-8-1 APU ECU PRI DC ESS BUS 2 SSPC 10A Logic: Always On
APU ECU SEC
Secondary 28 VDC
10 CDC 1
400 Hz on Ground CDC 5
CDC 1-5-16 APU SOV AUTO DC ESS SSPC 3A BUS 1 Logic: Combinational
DC EMER BUS
IASC 1 and IASC 2
APU FUEL SOV DC 3 EMER BUS
LGSCU
APU FIRE PANEL
EXTERNAL SERVICE PANEL
K6
AGCU Ready to Load
APU OVERHEAD SWITCH
Secondary Power Relay High
APU FIRE PBA Generator Load in
APU FEED SOV Close
APU Fuel SOV Command Close 28 VDC Primary Power APU Fuel SOV Command Open
DMC 1 and DMC 2
APU Shutdown
EXTERNAL TOWING SERVICE PANEL
APU FIRE SW
Pos Open
FIDEX CONTROL UNIT
28 VDC Return APU INLET DOOR
Excitation 28 VDC
APU COMPARTMENT SHUTDOWN SWITCH
LEGEND ARINC 429 AFDX
APU Shutdown
LEFT EEC and RIGHT EEC
CS1_CS3_4961_016
ADSP
Feedback
Figure 48: Electronic Control Unit Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
APU PROTECTIVE SHUTDOWNS The ECU provides protective, automatic shutdowns to prevent APU or aircraft damage. The weight-on-wheels (WOW) signal is used to inhibit some of the auto-shutdown functions when the APU is in flight mode. The ECU automatically shuts down the APU when certain fault conditions are detected. Protective shutdowns are indicated on the engine indication and crew alerting system (EICAS) by an APU SHUTDOWN advisory message with the exception of OVERSPEED, which is a dedicated warning message. In flight, the ECU inhibits some protective shutdown. When a protective shutdown is inhibited in flight, an APU caution message is displayed. The APU can be shut down manually if it is not required. During a protective shutdown, there is no cooling period and the APU shuts down immediately. When a protective shutdown occurs, the electrical and pneumatic loads are removed immediately. The bleed air valve closes and the SCV opens. Power is removed from the APU fuel control unit fuel solenoid and torque motor, and then the APU fuel SOV closes. The APU switch must be manually selected OFF when a protective shutdown occurs.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
Table 6: APU Protective Shutdown CONDITION
AUTO SHUTDOWN ON GROUND
AUTO SHUTDOWN IN FLIGHT
Failed/slow/hot/hung start
YES
YES
RPM underspeed
YES
NO
RPM overspeed/signal loss
YES
YES
Critical ECU fault (i.e. loss of DC power)
YES
YES
Loss of ARINC inputs to the ECU
YES
YES
APU fire
YES
NO
EGT overtemperature/signal loss
YES
NO
APU door position uncommanded/signal loss YES
NO
Excessive oil temperature
YES
NO
Low oil pressure
YES
NO
Oil filter clogged
YES
NO
APU inlet overheat/reverse flow
YES
NO
CS1_CS3_TB49_002
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the APU control system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
CAS MESSAGES Table 10: INFO Messages
Table 7: WARNING Message MESSAGE
LOGIC
APU OVERSPEED
MESSAGE
APU overspeed and no auto-shutdown.
49 APU FAULT - APU INOP
APU inoperative (APU fuel solenoid did not close when commanded fuel solenoid or loss of ARINC 429 input from DMC) or shutdown condition is detected.
49 APU FAULT - APU DEGRADED
Oil temp sensor fault or inlet temp sensor fault or ECU noncritical fault or APU low fuel flow and not inflight or oil pressure switch is lost.
49 APU FAULT - APU REDUND LOSS
APU secondary power relay command open short or APU 28 VDC secondary low or APU 28 VDC primary low and not APU 28 VDC secondary power relay command open short or APU 28 VDC secondary low and APU 28 VDC primary low or APU speed sensor failure and not loss of both speed sensors or EGT1 loss or EGT 2 loss and not EGT 1 and EGT 2 loss or APU ARINC 429 to DMC 1 no data received or APU ARINC DMC 2 no data received and not loss of both ARINC 429 signals.
Table 8: CAUTION Message MESSAGE
LOGIC
APU
APU failure in flight that does not lead to an automatic shutdown.
Table 9: ADVISORY Message MESSAGE
LOGIC
APU SHUTDOWN
Failures detected that cause the APU to shut down (mostly used on ground).
APU FAULT
Loss of redundant or noncritical function for the APU systems.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM The APU fuel shutoff valve is tested through the onboard maintenance system (OMS). The fuel shutoff valve test is found on the APU test page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-110
CS130_49_L2_APUControl_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-60 APU Control
MAINT MENU
RETURN TO LRU/SYS OPS
APU ECU-FUEL SHUT OFF VALVE TEST
001 /006 APU-FUEL SHUT OFF VALVE TEST PRECONDITIONS 1 - SET `APU RUN SW' TO OFF (APU SHUTDOWN) 2 - VERIFY THAT WEIGHT ON WHEELS IS TRUE
TEST INHIBIT/INTERLOCK CONDITIONS APU SHUTDOWN:
YYYYY
WEIGHT ON WHEELS:
YYYYY
PRESS `CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_4931_008
PRESS `CONTINUE' BUTTON TO GO TO NEXT PAGE
Figure 49: OMS Fuel Shutoff Valve Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-111
CS130_49_L2_APUControl_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
49-00 APU OPERATION GENERAL DESCRIPTION The auxiliary power unit (APU) is started from the flight deck using the APU switch located on the overhead panel. When the APU switch is momentarily held in the START position, the APU begins the start sequence: •
Built-in test (BIT)
•
APU door open
•
APU feed shutoff valve (SOV) open
When released, the selector automatically returns to RUN position. If the BIT tests are successful, the APU start is initiated. The electronic control unit (ECU) controls the APU start sequence. The ECU continuously monitors the APU start sequence and shuts down the APU if a fault is detected. If a BIT fails or a fault is detected during start, the APU switch is selected to OFF and a restart can be carried out after a two minute cooldown period. When the rotary selector is placed in the RUN position after START, the APU runs automatically, providing air and electrical loads as required. On ground, moving the APU switch to the RUN position without going to START enables the maintenance mode, which opens the APU inlet door and the aircraft APU feed shutoff valve. When the APU is running, selecting the APU switch to OFF initiates the cooldown sequence and the APU shuts down.
NOTE Moving the selector between OFF and RUN position requires a pull-to-turn action. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-112
CS130_49_L2_APUOperation_General.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU Switch Off
Start Sequence or Maintenance Mode
Start Sequence
APU Switch to START, then to RUN BIT Fail
Maintenance Mode
BIT Fail
APU Switch to RUN
BIT Starts BIT Intiated, Inlet Door and Bit Test APU Fuel SOV Open 5 Seconds BIT OK APU Switch to START, Release to RUN Select OFF Restart After 2 Minutes
APU Failed to Start
BIT OK
APU Start Sequence
Shutdown APU by Selecting Switch to OFF
APU Off
CS1_CS3_4961_017
APU Runing
Figure 50: APU Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-113
CS130_49_L2_APUOperation_General.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
CONTROLS AND INDICATIONS FLIGHT DECK CONTROLS APU Fire Panel The APU FIRE PBA is located on the ENGINE and APU FIRE panel on the overhead integrated cockpit control panel (ICCP). The APU FIRE PBA provides 28 VDC to close the APU feed shutoff valve through hardwired switch contacts. APU Panel The APU master switch is located on the left inboard ICCP module of the overhead panel. The switch position is sent to the ECU via an ARINC 429 BUS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-114
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU
APU FIRE
OFF
RUN
START
PULL TURN
BTL
APU PANEL
CS1_CS3_4961_011
ENGINE AND APU FIRE PANEL
Figure 51: APU Flight Deck Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-115
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
EXTERNAL CONTROLS There are two external APU shutdown switches. Operating either switch causes the APU to shut down without a cooldown period. The switch inputs are inhibited when the aircraft is in flight. ELECTRICAL/TOWING SERVICE Panel The APU SHUT OFF toggle switch on the external service panel is a guarded momentary contact switch with a center OFF position. The switch is hardwired to the data concentrator unit module cabinet (DMC). APU Compartment Shutdown Switch The APU compartment shutdown switch is mounted on the forward firewall. The switch is hardwired to the APU ECU.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-116
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
MAINT LTS ON/OFF
NAV LTS
Shutdown Switch
APU SHUT OFF
LAMP TEST
CKPT IN USE EXT PWR
TOWBARLESS ONLY
AVAIL
NO TOW
IN USE
TOW
SERV
BATT
ON TOW PWR
OMS/HMU OMS/ MS/HMU
C CALL
ON
HEADSET
APU COMPARTMENT SHUTDOWN SWITCH
CS1_CS3_4961_013
ELECTRICAL/TOWING SERVICE PANEL
Figure 52: APU External Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-117
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
SYNOPTIC PAGE The APU performance information is displayed on the STATUS synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-118
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
STATUS
AIR IR
DOOR OR
ELEC C
FLT CTR CTRL L
FUEL FUE EL
HYD YD
AVIONIC IONIC
INFO
CB
15 °C 15 °C
TAT SAT
APU RPM Symbol
ENGINE
103 113 21.8
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L) APU
RPM EGT DOOR
100 % 312 °C OPEN
OIL TEMP OIL PRESS OIL QTY
85
APU operation or below redline.
107
APU operation above redline (overspeed).
è½½
APU rpm value invalid.
75 °C NORM NORM APU EGT
TIRE PRESSURE (PSI) Symbol
150 224
150
224
224
Condition
224
Condition
650
APU operating at or below redline.
960
APU operating above redline (overtemperature).
è½½
APU EGT invalid.
BRAKE
00
TEMP
01
01
CS1_CS3_4991_011
00
SYNOPTIC PAGE - STATUS
Figure 53: APU Status Performance Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-119
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
•
OPERATION APU START On the APU control panel, pull and turn the APU switch to the START position and release to RUN. •
APU inlet door opens
•
On the STATUS synoptic page:
•
-
APU DOOR indicates open
-
APU RPM increases
-
RPM 91 to 100% (ground) 100% (flight)
-
EGT less than 950°C
-
DOOR OPEN
-
OIL TEMP NORM
-
OIL PRESS NORM
-
OIL QTY FULL
WARNING
On EICAS: -
•
On STATUS synoptic page:
APU IN START status message
On the FUEL synoptic page: -
APU feed SOV open
-
APU flow line green
-
APU fuel filter green
BEFORE OPERATING THE APU, MAKE SURE THAT THE ACCESS DOORS ARE LOCKED IN THE OPEN OR CLOSED POSITION. DO NOT OPEN OR CLOSE THE ACCESS DOORS DURING APU OPERATION. THE APU AIR SUCTION CAN CAUSE THE DOORS TO CLOSE QUICKLY IF THEY ARE NOT LOCKED AND CAUSE INJURY TO PERSONS AND DAMAGE TO THE EQUIPMENT.
NOTE APU IN START is displayed until APU reaches 70% rpm. At 70% rpm, the following message is shown on EICAS: •
APU ON status message
At 100% rpm: •
On EICAS -
APU ON status message
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
CAUTION 1. The APU start from battery shall only be allowed if there are two batteries available and TRU 3 is not available. A battery start is allowed with one battery available and a TRU feeding the DC BUSES when TRU 3 is not available. 2. The APU fire extinguishing system is not effective when the access doors are open. If the APU is operated with the access doors open, make sure that fire fighting equipment is available.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-120
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU RUN
OFF
APU
NOTE
START
OFF
Do not perform more than three starts/start attempts in one hour. A two-minute delay must be observed between start attempts.
PU TU LL RN
RUN
START
PULL TURN
APU PANEL
APU PANEL
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
4000 KG 0 KG
20 °C 2000 KG
20 °C 2000 KG
APU IN START
APU ON
APU ON 0 KG
EICAS
EICAS
EICAS
APU
EGT DOOR
0% 22 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
25 °C NORM FULL
RPM EGT DOOR
SYNOPTIC PAGE – STATUS
SYNOPTIC PAGE – FUEL
70 % 285 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
32 °C NORM NORM
RPM EGT DOOR
100 % 315 °C OPEN
OIL TEMP OIL PRESS OIL QTY
97 °C NORM NORM
SYNOPTIC PAGE – STATUS
SYNOPTIC PAGE – STATUS
70% rpm
100% rpm
0% rpm
CS1_CS3_4961_019
APU RPM
Figure 54: APU Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-121
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
NORMAL APU SHUTDOWN The normal shutdown is carried out using the APU switch in the flight deck. Before shutting down, remove the electrical and pneumatic loads. On the APU control panel, pull and turn the APU switch to the OFF position. After one minute: •
APU stops, RPM decreases to 0%, and EGT decreases towards ambient temperature
•
On the FUEL synoptic page, the APU SOV shows CLOSED, and the APU flow line shows white
•
On the STATUS page DOOR CLOSED is shown in white
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-122
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU OFF
APU
RUN
START
OFF
RUN
LL PU N R TU
START
LL PU N R TU
60 Second Time Delay APU PANEL
APU PANEL
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
4000 KG 0 KG
60 Second Time Delay Ends
20 °C 2000 KG
20 °C 2000 KG
APU ON EICAS
0 KG
EICAS APU
EGT DOOR
100 % 315 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
97 °C NORM NORM
SYNOPTIC PAGE – STATUS
RPM EGT DOOR
0% 125 °C CLOSED
OIL TEMP OIL PRESS OIL QTY
32 °C NORM FULL
SYNOPTIC PAGE – STATUS
100% RPM
SYNOPTIC PAGE – FUEL 0% RPM
CS1_CS3_4961_020
APU RPM
Figure 55: APU Normal Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-123
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
EMERGENCY APU SHUTDOWN The emergency shutdown is carried out from the external shut down switches or the APU FIRE PBA in the flight deck. Pressing any emergency shutdown switch: •
Immediately APU stops with no cooldown period, RPM decreases to 0%, and EGT decreases toward ambient temperature
•
On FUEL synoptic page, APU SOV shows CLOSED and APU flow line shows white
•
On STATUS page DOOR CLOSED is shown in white
On the APU control panel, pull and turn the APU switch to the OFF position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-124
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU
APU
APU OFF
RUN
APU
START
OFF
FIRE
RUN
START
FIRE PULL TURN
BTL
PULL TURN
BTL
AVAIL
AVAIL
APU PANEL
APU PANEL ENGINE AND APU FIRE PANEL
ENGINE AND APU FIRE PANEL
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
4000 KG 0 KG
No Cooldown Period
20 °C 2000 KG
20 °C 2000 KG
APU ON EICAS
0 KG
EICAS APU
EGT DOOR
100 % 315 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
97 °C NORM NORM
SYNOPTIC PAGE – STATUS
RPM EGT DOOR
0% 125 °C CLOSED
OIL TEMP OIL PRESS OIL QTY
32 °C NORM FULL
SYNOPTIC PAGE – STATUS
100% RPM
SYNOPTIC PAGE – FUEL 0% RPM
CS1_CS3_4961_021
APU RPM
Figure 56: APU Emergency Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-125
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
MONITORING AND TEST The following page provides the crew alerting system (CAS) messages for the APU operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-126
CS130_49_L2_APUOperation_MonitoringTest.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
CAS MESSAGES Table 11: STATUS Messages MESSAGE
LOGIC
APU ON
APU is running.
APU IN START
APU performing startup sequence.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-127
CS130_49_L2_APUOperation_MonitoringTest.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-128
CS130_49_ILB.fm
ATA 71 - Power Plant
BD-500-1A10 BD-500-1A11
71 - Engine
Table of Contents 71-10 Nacelles........................................................................71-2 General Description ..........................................................71-2 71-60 Inlet Cowl .....................................................................71-4 General Description ..........................................................71-4 71-11 Fan Cowls ....................................................................71-6 General Description ..........................................................71-6 Practical Aspects ..............................................................71-8 78-30 Thrust Reverser Doors ...............................................71-12 General Description ........................................................71-12 Component Location ......................................................71-24 Component Information ..................................................71-26 Detailed Description ........................................................71-30 Practical Aspects ............................................................71-32 78-11 Turbine Exhaust System ............................................71-38 General Description ........................................................71-38 71-40 Engine Mounts and Pylon...........................................71-40 General Description ........................................................71-40 Component Location ......................................................71-40 Detailed Component Information ....................................71-42 Practical Aspects ............................................................71-46 71-70 Engine Drains .............................................................71-54 General Description ........................................................71-54 71-00 Engine Storage and Preservation..............................71-56 General Description ........................................................71-56
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-i
CS130_M5_71_PowerPlantTOC.fm
71 - Engine
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7: Figure 8: Figure 9: Figure 10: Figure 11: Figure 12:
Nacelle Systems....................................................71-3 Inlet Cowl...............................................................71-5 Fan Cowls .............................................................71-7 Fan Cowl Opening and Closing.............................71-9 Fan Cowl Latch and Keeper Adjustment.............71-11 Thrust Reverser Doors ........................................71-13 Thrust Reverser Fixed Structure .........................71-15 Thrust Reverser Door Latches ............................71-17 Bumpers and Bifurcation Latch System ..............71-19 Power Door Operating System............................71-21 Power Door Operating System Control ...............71-23 Power Door Operating System Component Location ...........................................71-25 Figure 13: Power Door Operating System Powerpack...........................................................71-27 Figure 14: Power Door Operating System Locking Actuator and Hold Open Rods ...............71-29 Figure 15: Power Door Operating System Schematic..........71-31 Figure 16: Thrust Reverser Door Bifurcation Latch System and Latch Operation .....................71-33 Figure 17: Thrust Reverser Door Opening and Closing .......71-35 Figure 18: Thrust Reverser Door Bifurcation Latch System Adjustment....................................71-37 Figure 19: Turbine Exhaust System .....................................71-39 Figure 20: Engine Mounts and Pylon ...................................71-41 Figure 21: Forward Engine Mount ........................................71-43 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 22: Aft Engine Mount and Thrust Links..................... 71-45 Figure 23: Engine Removal Bootstrap Equipment............... 71-47 Figure 24: Engine Cradle, and Transport Stand .................. 71-49 Figure 25: Electrical Harness Disconnects .......................... 71-51 Figure 26: Hydraulic and Fuel Connections......................... 71-53 Figure 27: Engine Drains ..................................................... 71-55 Figure 28: Engine Storage and Preservation....................... 71-57 Figure 29: Engine Preservation Chart.................................. 71-59
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-ii
CS130_M5_71_PowerPlantLOF.fm
71 - Engine
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-iii
CS130_M5_71_PowerPlantLOF.fm
POWER PLANT - CHAPTER BREAKDOWN POWER PLANT CHAPTER BREAKDOWN
Nacelle Systems
1
Power Door Operating System
2
Engine Mounts and Pylon
3
Engine Storage - Preservation
4
71 - Power Plant 71-10 Nacelles
71-10 NACELLES GENERAL DESCRIPTION The engine cowlings minimize airflow turbulence around the engine. They provide an aerodynamically smooth surface around the engine and protect the engine mounted components. The engine inlet cowl assembly directs the airflow to the engine fan. The inlet cowl is acoustically treated for noise reduction. The fan cowls, installed between the inlet cowl and the thrust reverser doors, are attached to the pylon fan cowl support beam. Each inboard fan cowl has a strake installed to decrease the turbulence caused by the airflow between the fan cowl and the fuselage. The thrust reverser doors enclose the core engine and contain the necessary components for the thrust reverser system. The thrust reverser doors attach to the hinge beam on the pylon. Both the fan cowls and thrust reverser doors can be opened to provide access to the engine. The turbine exhaust system controls and directs the exhaust gases.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-2
CS130_71_L2_Nacelles_General.fm
71 - Power Plant 71-10 Nacelles
Strake
Pylon
Turbine Exhaust System
Fan Cowl
Thrust Reverser Door CS1_CS3_7100_015
Inlet Cowl
Figure 1: Nacelle Systems Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-3
CS130_71_L2_Nacelles_General.fm
71 - Power Plant 71-60 Inlet Cowl
71-60 INLET COWL GENERAL DESCRIPTION The inlet cowl provides smooth undisturbed airflow to the fan in all operating conditions, and reduced noise. The inner barrel of the inlet cowl is made of titanium and composite with 360° of interior acoustic treatment. Lightning strike protection is provided by a copper screen layer impregnated into the outer barrel. The outer barrel is a two-piece composite structure extending from the inlet lip skin to the leading edge of the fan cowl, providing a smooth surface for airflow across the nacelle exterior. Two hoist points are provided for the removal of the inlet cowl. A NACA scoop at the 1 o’clock position is used for fan compartment cooling. A panel, located at the 11 o’clock position, provides access to the P2T2 probe, sense line, and wiring harness. The inlet cowl is anti-iced by bleed air. The air is exhausted out of the bottom of the inlet cowl through a series of slots. An access panel for the cowl anti-ice (CAI) duct is located at the 5 o’clock position. Drain holes in the core of the inner barrel and at the bottom of the outer barrel allow fluids to drain. The inlet cowl is bolted to the front flange of the engine at the attachment ring.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-4
CS130_71_L2_InletCowl_General.fm
71 - Power Plant 71-60 Inlet Cowl
P2T2 Probe
Hoist Points
NACA Scoop
Inner Barrel Sound Proofing
P2T2 Access Panel
Outer Barrel
Attachment Ring Inlet Cowl Sling
Anti-ice Air Exhaust Louvers
Cowl Anti-Ice Duct Access Panel
Lip
CS1_CS3_7111_002
Drain Holes
Figure 2: Inlet Cowl Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-5
CS130_71_L2_InletCowl_General.fm
71 - Power Plant 71-11 Fan Cowls
71-11 FAN COWLS GENERAL DESCRIPTION The fan cowls cover the fan case, providing a protective enclosure for engine mounted components and a smooth aerodynamic outer surface for the nacelle. They are hinged at the top and latched at the bottom centerline with three flush mounted latches. The fan cowl doors are manufactured using carbon fiber reinforced polymer (CFRP), strengthened with radial stiffeners and longerons. Copper mesh is embedded in the fan cowl laminate for lightning strike protection. The removable strakes are installed on the inner cowls only. A pressure-relief door, located on the outer fan cowls, relieves high-pressure in the event of a duct failure. During maintenance, hold open rods (HORs) provide a means to support the cowls in the open position. Drain holes, located near the lower centerline, ensure fluid drainage. Three hoist points of two screw holes are on each panel. They are used to attach the fan cowl door assembly to the ground handling sling. The ground handling sling is used to lift the cowl during removal or installation. Wear strips ensures aerodynamic smoothness is maintained between the cowls. Three hinges attach the cowls to the pylon. These hinges are secured by pins held in place by retaining clips. The pins facilitate removal of the fan cowl from the pylon. Two HORs on each panel are used to hold the cowls in the open position. Quick-release fittings secure one end of the rods to the fan case when the cowls are opened. The other end of the rods are attached to a pivot bracket assembly. Stow brackets and clips on the cowl provide a means to secure the rods when the cowl is closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-6
CS130_71_L2_FanCowls_General.fm
71 - Power Plant 71-11 Fan Cowls
Stiffeners
Quick-Release Fitting Hinges Hold Open Rod Clip
Pressure-Relief Door
Strake
Wear Strips
Hoist Attachment Points
Latches
HOLD OPEN ROD
CS1_CS3_7111_004
Drain Holes
Figure 3: Fan Cowls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-7
CS130_71_L2_FanCowls_General.fm
71 - Power Plant 71-11 Fan Cowls
PRACTICAL ASPECTS
•
Always refer to the latest Aircraft Maintenance Publication (AMP) before carrying out any procedures on the aircraft.
Put the forward HOR on to the HOR clip. Secure the HOR with the snap spring of the clip
•
Carefully lower the fan cowl and push the fan cowl door together at the bottom centerline. Make sure the alignment pins enter the holes adjacent to the latches
•
Engage the hooks on all three latches with the eye-bolts on the three latch keepers in the sequence L1, L3, L2
•
Close the three latch handles until they are flush with the door surface and locked in position. When the latches are properly closed, the red marks on the latches do not show
OPEN THE FAN COWL To open the fan cowl: •
Push the fan cowl door latch triggers to release the three latch handles. Pull down each handle to open the three latch handles
•
Move the latches away from the three latch keepers in the sequence L1, L3, and L2
•
Manually lift and hold a fan cowl door at the lower edge
•
Remove the forward hold open rod (HOR) from the fan cowl and install it in the forward HOR bracket on the fan case. Secure it with the quick-release pin on the end of the HOR
•
Remove the aft HOR from the fan case and secure it to the power door operating system (PDOS) bracket on the fan case. Secure it with the quick-release pin on the end of the HOR
•
Slowly lower the fan cowl door until the HORs hold the weight of the door
CLOSE THE FAN COWL To close the fan cowl: •
Manually lift and hold the fan cowl door at the lower edge so that the weight is not on the HORs
•
Remove the aft HOR from the PDOS bracket on the fan case and install it on the fan cowl. Secure the HOR with the quick-release pin
•
Remove the forward HOR from the forward HOR bracket on the fan case and install it on the fan cowl. Secure the HOR with the quick-release pin
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
WARNING 1. TWO OR MORE PEOPLE ARE REQUIRED TO OPEN THE FAN COWL DOOR. THE FAN COWL DOOR WEIGHS MORE THAN 45.36 KG (100 LB). IF THE DOOR FALLS, INJURY AND/OR DAMAGE TO THE ENGINE OR PERSON(S) CAN OCCUR. 2. DO NOT KEEP OPEN A FAN COWL DOOR WHEN THE WIND SPEED IS OVER 96 KM/H (60 MPH). IF THE WIND MOVES THE FAN COWL DOOR, INJURY AND/OR DAMAGE TO THE ENGINE OR PERSON(S) CAN OCCUR. 3. CAUTION IS REQUIRED IF YOU OPEN A FAN COWL DOOR WHEN THE WIND SPEED IS 74 KM/H (46 MPH) OR MORE. IF THE WIND MOVES THE FAN COWL DOOR, INJURY AND/OR DAMAGE TO THE ENGINE OR PERSON(S) CAN OCCUR.
CAUTION Lifting the fan cowl door more than 55° vertical can cause fan cowl door damage or pylon damage.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-8
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 71-11 Fan Cowls
A
LATCH SEQUENCE OPENING: L1, L3, L2 CLOSING: L1, L3, L2
FAN COWL OPEN (VIEW LOOKING FORWARD)
CS1_CS3_7111_010
C05210065-001
A
Figure 4: Fan Cowl Opening and Closing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-9
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 71-11 Fan Cowls
FAN COWL LATCH AND KEEPER ADJUSTMENT The fan cowl latches require adjustment anytime the fan cowl does not close properly and the latches are safely engaged. The latches engage adjustable latch keepers on the opposite cowl. If the fan cowl latches require adjustment, the keepers are adjusted by inserting an allen key into the star wheel adjustment nut. The latches are adjusted to provide a closing force of 9.1 to 10.9 kgf (20 to 24 lbf). Shim plates on latches and keepers are used to adjust the cowl gap. Make sure the gap between the left and right fan cowls is 0.51 - 5.08 mm (0.020 - 0.200 in.).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-10
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 71-11 Fan Cowls
Shim Plates
Fan Cowl Latch Keepers B
C
A A Fan Cowl Latches
Star Wheel Adjuster
Keeper Hook
Latch
Star Wheel Adjuster
Allen Key
C
B
CS1_CS3_7111_003
Latch Handle
Figure 5: Fan Cowl Latch and Keeper Adjustment Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-11
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
78-30 THRUST REVERSER DOORS GENERAL DESCRIPTION The power plant is equipped with a cascade type thrust reverser (T/R) assembly consisting of left and right T/R doors. The left and right T/R doors are installed on a hinge beam on the pylon. The hinge beam has four attachment points. The T/R doors are latched together at the bottom of the engine and form an enclosure for the core engine. The T/R doors can be opened using the power door operating system (PDOS), or manually using a hydraulic pump. When opened, the engine core can be accessed. Each T/R door consists of two fixed structures that direct fan air aft. The inner structure forms an aerodynamic surface from the fan case exit to the aft nozzle. The outer structure contains the T/R translating sleeve, cascades, and blocker doors that provide reverse thrust. The inner and outer fixed structures are connected by the torque box. The torque box is the attachment point for the T/R components and the PDOS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-12
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Hinge Beam
Blocker Door
Cascade Segments
Torque Box
Inner Fixed Structure Latch Beam
CS1_CS3_7110_001
Translating Sleeve
Figure 6: Thrust Reverser Doors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-13
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
THRUST REVERSER FIXED STRUCTURE The thrust reverser fixed structure covers the engine core. It provides core compartment ventilation, and forms the fan duct inner aerodynamic surface from the fan case exit to the exhaust nozzle. The thrust reverser attaches to the pylon at the hinge beam. An air/oil cooler (AOC) inlet and exhaust is located on the left side of the fixed structure. A pressure-relief door, located behind the AOC, opens in the event of a burst duct in the bleed air system. The pressure relief door is located on the left side of the inner fixed structure behind the air/oil cooler (AOC) exhaust, at the 9 o'clock position. The latches release at 3.5 psi. The door is manually closed if it has opened. The right side of the fixed structure has an oil tank service door and a cutout for the active clearance control (ACC) ducting. A precooler exhaust door is located behind the oil tank service door. Forward and aft bumpers, located at the top and bottom of the structure, are used to align the doors as they are closing. A locking latch is used to ensure the doors remain closed in the aligned position. Latches, installed on the latch beam, are used to secure the T/R door in the closed position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-14
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
AOC Inlet and Exhaust
Hinge Beam
Fwd Upper Bumper
ACC Duct
Aft Upper Bumper
Oil Tank Service Door Precooler Exhaust Door
Locking Latch
Aft Upper Bumper
Fwd Lower Bumper Pressure-Relief Door Latch Beam Aft Latch
Aft Lower Bumper
Aft Lower Bumper
LEFT THRUST REVERSER FIXED STRUCTURE
RIGHT THRUST REVERSER FIXED STRUCTURE
CS1_CS3_7111_007
LEGEND ACC AOC
Aft Latch
Active Clearance Control Air/Oil Cooler
Figure 7: Thrust Reverser Fixed Structure Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-15
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
THRUST REVERSER DOOR LATCHES Latches are quick-release mechanisms integral to the latch beam that provide quick access to the engine core. The latches provide resistance to loads that might otherwise cause the T/R to disengage or open during the flight cycle. Latches no. 1, no. 2, no. 3, and no. 4 are installed on the latch beam. Latch no. 5 is located on the lower aft section of the inner fixed structure. The latches are takeup latches that pull the two T/R halves together when securing. Latches no. 2 and no. 3 are located inside the latch access door. The latch access door cannot be closed if latch 2 and latch 3 are not properly secured. A closure assist assembly consists of a turnbuckle, which is used to draw the door together before engaging the latches. An eye on one end of the turnbuckle allows it to pivot in order for the pin to engage the opposite cowling. When engaged, the turnbuckle is turned manually to draw the two doors together. The closure assist assembly is stowed after use.
CAUTION 1. Do not turn the body of the closure assist assembly to make it shorter than the minimum length. The closure assist assembly can be damaged if the minimum length is decreased. 2. Adjust the closure assist assembly to the minimum length before placing the retention clip of the stow bracket assembly. The closure assist assembly can move out of position if this step is not followed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-16
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Closure Assist Assembly Stowed Position
Latch No. 2
A
Latch No. 3
Latch No. 5
Latch No. 1 Latch No. 4
Latch Access Door
CS1_CS3_7111_009
A
Figure 8: Thrust Reverser Door Latches Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-17
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Bumpers and Bifurcation Latch System To maintain a correct fit of the ducts to the engine, the inner fixed structure (IFS) has upper and lower bumpers to assist alignment as the two thrust reverser doors are closed. The upper bumper includes a compression strut to compensate for the larger gap between the two cowls. The bifurcation latch system (BLS) is similar to the bumpers but has a locking feature to keep it engaged. The handle is accessed through the latch access door. The BLS is used to limit the deflection of the IFS, and maintain its structural integrity in the event of a burst air duct. When engaged, the bifurcation latch uses a rod and cam mechanism to insert a locking pin through the latch. As the thrust reverser (T/R) doors close, a pin on the left side is inserted into a receiver located on the right side. The pin mechanism is locked together using a rod. The rod is inserted through the pin by turning, and pushing the BLS lever up. The bifurcation latch lever is painted red, providing a visual indication of the position. If the bifurcation latch lever is not locked, the latch access panel cannot be closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-18
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Compression Bumper
Bumpers Latch Fitting
Adjustable Pin
Upper Bumpers
Compression Strut
Rod
Latch Beam Attach
Bifurcation Latch System Handle
CS1_CS3_7100_004
Lower Bumpers
Figure 9: Bumpers and Bifurcation Latch System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-19
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
POWER DOOR OPERATING SYSTEM The thrust reverser (T/R) power door operating system (PDOS) is used to open the T/R doors and support them in the open position. Each T/R door is opened independently of the other. The PDOS powerpack consists of a hydraulic gear pump driven by an AC motor, and a reservoir capable of supplying fluid to two locking actuators. The PDOS is operated by switches located on each side of the fan case. Each T/R door is raised and lowered using a locking actuator connected between the engine fan case and the T/R door torque box. The locking actuator has an internal locking mechanism that is engaged when the actuator is fully extended. A hold open rod (HOR) provides additional support for the T/R door. The locking actuator takes 100% of the load under normal operation, with the HOR as a failsafe mechanism. Each locking actuator is connected to the hydraulic powerpack using a hydraulic line with a quick-disconnect fitting. If the powerpack fails, the quick-disconnect fitting can be removed from the locking actuator and a hand pump connected in its place.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-20
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Locking Actuator
Powerpack
Hold Open Rod Power Door Operating System Control Switch
CS1 CS3 7100 013 CS1_CS3_7100_013
Hydraulic Line
Figure 10: Power Door Operating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-21
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Power Door Operating System Control The PDOS can be operated when the aircraft is on the ground and the engine is shut down. The left engine powerpack motor is powered by 115 VAC from AC BUS 1. The PDOS operates when the thrust reverser (T/R) door switch is pressed. Only one door can be opened at a time. The switch supplies 28 VDC from DC BUS 1 to energize solenoids within the powerpack. When the solenoid is energized, the powerpack supplies 2300 psi hydraulic pressure to the locking actuator. The right engine PDOS motor is powered from AC BUS 2, and the control power is supplied from DC BUS 2. When the actuator fully extends, a pressure-relief valve (PRV) in the actuator opens to ensure the maximum system pressure is not exceeded. The T/R door switch is held for 5 seconds after full actuator travel, then released. The PRV closes, and the internal lock engages and holds the door open. To close the door, the T/R door switch is held. The actuator extends slightly and the internal lock releases. When the lock is released, the switch is released. The actuator retracts due to the weight of the T/R door. An internal restrictor limits the rate of closing. If no power is available, a hand pump can be connected to the actuator quick-disconnect fitting to open the door.
CAUTION 1. The inboard thrust reverser doors can contact the slats if the slats are in the extended position. 2. The actuators can be opened manually using a hand pump. The introduction of additional fluid in the system may cause the powerpack reservoir to overflow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-22
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
DC BUS 1
Left Thrust Reverser Door Switch
Right Thrust Reverser Door Switch
POWERPACK
Left Thrust Reverser Door Actuator
Right Thrust Reverser Door Actuator
Hand Pump NOTE LEGEND
Left engine shown. Right engine similar.
Hydraulic Pressure
CS1_CS3_7100_016
AC BUS 1
Figure 11: Power Door Operating System Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-23
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
COMPONENT LOCATION The power door operating system (PDOS) consists of the following components: •
Powerpack
•
Locking actuators
•
Control switches
•
Hold open rods (HORs)
POWERPACK The powerpack is located on the fan case at the 10 o’clock position. LOCKING ACTUATORS The locking actuators are located on the fan case at the 3 o’clock and 9 o’clock positions. CONTROL SWITCHES The control switches are located on the fan case at the 4 o’clock and 8 o’clock positions. HOLD OPEN RODS The hold open rods (HORs) are installed on each thrust reverser (T/R) door torque box.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-24
CS130_71_L2_TRD_ComponentLocation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Powerpack
Hold Open Rod Power Door Operating System Control Switch
CS1_CS3_7100_006
Locking Actuator
Figure 12: Power Door Operating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-25
CS130_71_L2_TRD_ComponentLocation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
COMPONENT INFORMATION POWERPACK The self-contained power door operating system (PDOS) is powered by an electrohydraulic powerpack producing 2300 psi. The powerpack houses an AC motor and a gear pump assembly, which provides hydraulic pressure to two actuators during system operation. The powerpack includes a reservoir, solenoid valves, filter, and a filler cap/strainer. A drain/vent port on the lower half of the powerpack provides venting and overflow drainage. The system is serviced with engine oil. A filler cap on the reservoir lid is used to replenish the oil. A dipstick is provided under the filler cap for checking the oil level. A strainer prevents any contaminants from entering the reservoir. Oil levels are marked on the dipstick as follows: •
L - Low 0.74 L
•
N - Normal 0.82 L
•
H - High 0.89 L
CAUTION Before servicing the PDOS powerpack, the thrust reverser (T/R) doors should be fully closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-26
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Filler Cap/Strainer and Dipstick
Solenoid Valves Drain/Vent Port
CS1_CS3_7100_007
AC Motor
Figure 13: Power Door Operating System Powerpack Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-27
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
LOCKING ACTUATORS Two locking actuators are used independently to open each thrust reverser (T/R) door. Each actuator has a flow control valve (FCV), quick-disconnect connector, pressure-relief valve (PRV), mechanical lock, and a manual bleed feature. The actuators internal flow control valve provides the system with the means to control the closure rate of the cowl when the lock is released. Both actuators have quick-disconnect fittings to accommodate a hand pump in the event of a powerpack failure. When opened, an automatic mechanical lock inside the actuator provides positive retention of the C-duct. A manual bleed port on the actuator is used to expel air from the system. HOLD OPEN RODS Each T/R door has a telescopic hold open rod (HOR) which is attached between the T/R door and the fan case to help hold the T/R doors open. Lock collars are used to extend and retract the HOR. The HOR is used in conjunction with the locking actuator to hold the T/R door open, and provide a safe environment during maintenance. When not in use, the HORs are stored on the T/R door torque box.
WARNING 1. MAKE SURE THAT THE HOLD OPEN RODS (HORS) ARE INSTALLED. FAILURE OF THE POWER DOOR OPERATING SYSTEM (PDOS) CAN CAUSE THE THRUST REVERSER (T/R) TO CLOSE IF THE HORS ARE NOT INSTALLED. THIS CAN CAUSE INJURY TO PERSONNEL OR DAMAGE TO EQUIPMENT.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-28
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Mechanical Lock Pin Rod End
Manual Bleed Port
Cylinder
A
Bearing Quick-Disconnect Fitting A LOCKING ACTUATOR (RETRACTED)
Lock/Release Collar
Unlock
HYDRAULIC HAND PUMP
Lock HOLD OPEN ROD
CS1_CS3_7100_008
Extend/Retract Collar
Figure 14: Power Door Operating System Locking Actuator and Hold Open Rods Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-29
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
DETAILED DESCRIPTION The power door operating system (PDOS) is powered when the aircraft is on the ground and the engine is shut down. For the left engine, the following logic is supplied to the control and distribution cabinet 1 (CDC 1). •
Weight-on-wheels (WOW) signal from the landing gear and steering control units (LGSCUs)
•
Engine shutdown signal from the electronic engine control (EEC)
When the thrust reverser (T/R) door switch is pressed, the solenoid valve is energized, allowing hydraulic fluid to be supplied to the hydraulic pump. When solenoid is energized, logic turns on the L (R) PDOS Motor SSPC and the motor drives the hydraulic pump. The thrust reverser door hydraulic actuator extends to open the door. If the motor should overheat, a thermal switch shuts the motor off.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-30
CS130_71_L3_TRD_DetailedDescription.fm
71 - Power Plant 78-30 Thrust Reverser Doors
ELECTRONIC ENGINE CONTROL
LEFT DOOR REVERSER DOOR SWITCH
DMC 1 AND DMC 2
PDOS POWERPACK
SOL
LGSCU 1 AND LGSCU 2
CDC 1
WOW L ENG OFF
to Left Thrust Reverser Door Actuator
CDC 1-11-12 L PDOS Ctrl DC SSPC 3A BUS 1
Hydraulic Reservoir
Hydraulic Pump to Right Thrust Reverser Door Actuator
Logic: Combinational
L ENG OFF L or R Solenoid Energized
CDC 1-17-4-5-6 L PDOS Motor AC BUS 1 SSPC 7.5A
SOL
WOW
Motor
Logic: Combinational Thermal Switch
LEGEND
SOL
Hydraulic Pressure Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Landing Gear and Steering Control Unit Power Door Operating System Weight-on-Wheels AFDX ARINC 429 Solenoid
RIGHT DOOR REVERSER DOOR SWITCH
NOTE - Left engine shown, right engine similar. - Left thrust reverser door shown operating.
CS1_CS3_7100_009
CDC DMC LGSCU PDOS WOW
Figure 15: Power Door Operating System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-31
CS130_71_L3_TRD_DetailedDescription.fm
71 - Power Plant 78-30 Thrust Reverser Doors
PRACTICAL ASPECTS THRUST REVERSER DOOR BIFURCATION LATCH SYSTEM AND LATCH OPERATION Open the Bifurication Latch System 1. Open the latch access door.
CAUTION The closure assist assembly must be at its minimum length before being put into the retention clip of the stow bracket assembly. If this is not done, the closure assist assembly can move out of position.
2. Turn the bifurication latch system (BLS) latch handle 90° clockwise.
Close the Thrust Reverser Door Latches
3. Pull the BLS latch handle down until it comes to a stop.
1. Make sure that the T/R doors are fully closed.
4. Turn the BLS latch handle 90° counterclockwise.
2. Adjust the closure assist assembly so that it can engage the hook on the fixed panel support on the other T/R half.
Open the Thrust Reverser Door Latches 1. Remove the closure assist assembly from the stow bracket assembly and adjust it to its minimum length. 2. Engage the closure assist assembly in the hook on the fixed panel support on the other thrust reverser (T/R) door.
3. Turn the body of the closure assist assembly with a wrench to pull the two T/R halves closer together. 4. Close the T/R door latches in the sequence that follows: -
LATCH No. 1
3. Turn the body of the closure assist assembly with a wrench to pull the two T/R doors together and relieve the tension on the T/R latches.
-
LATCH No. 2
-
LATCH No. 3
4. Open the T/R door latches in the sequence that follows.
-
LATCH No. 4
-
LATCH No. 5
-
LATCH No. 5
-
LATCH No. 4
-
LATCH No. 3
-
LATCH No. 2
-
LATCH No. 1
5. Put the closure assist assembly into its stowed position Close the Bifurication Latch System 1. Turn the BLS latch handle 90°clockwise.
5. Put the closure assist assembly into its stowed position. Ensure the closure assist assembly is at its minimum length.
2. Push the BLS latch handle up until it comes to a stop. 3. Turn the BLS latch handle 90° counterclockwise. 4. Close the latch access door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-32
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Latch No. 3
Latch Access Door
Latch No. 5
Latch No. 4
Latch No. 2
Closure Assist Assembly Stowed Position
Closure Assist Assembly Engaged Position
Bifurcation Latch System Handle
CS1_CS3_7100_017
Latch No. 1
Figure 16: Thrust Reverser Door Bifurcation Latch System and Latch Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-33
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
THRUST REVERSER DOOR OPENING AND CLOSING Open the Thrust Reverse Door
Close the Thrust Reverser Door
1. Push PDOS operating switch until the locking actuator is fully extended.
1. Release the HOR from the HOR open fitting on the fan case and install the HOR in the HOR stow bracket.
1. Hold the switch for a minimum of 5 seconds after the locking actuator is fully extended and the powerpack pressure-relief valve (PRV) opens.
2. Push the PDOS operating switch until the locking actuator is fully extended.
2. Release the switch and the locking actuator retracts in a controlled descent until the mechanical lock is engaged. 3. Release the top end of the hold open rod (HOR) from the HOR stow bracket and install it in the HOR open fitting on the fan case.
WARNING 1. DO NOT MOVE BETWEEN THE ENGINE AND THE OPEN THRUST REVERSER (T/R) UNTIL THE THE HOLD OPEN ROD (HOR) IS INSTALLED. THE T/R IS HEAVY AND CAN CLOSE QUICKLY IF THE PDOS FAILS. 2. TO PREVENT INJURY TO PERSONNEL OR DAMAGE TO EQUIPMENT, CAUTION IS REQUIRED WHEN OPENING THE T/R IN HIGH WIND.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
3. Hold the switch for a minimum of 5 seconds after the locking actuator is fully extended and the powerpack PRV opens. 4. Release the switch. The powerpack PRV closes and the locking actuator mechanical lock disengages. 5. Make sure that the thrust reverser outer V-groove blade engages with the outer V-groove on the engine during closing. 6. Make sure that the alignment pins correctly engage the receptacles before the thrust reverser doors are fully closed.
CAUTION Make sure that the T/R half area is clear of tools and equipment before closing the door. Failure to do so can cause damage to the T/R half and to the engine.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-34
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Hold Open Rod Stow Bracket
Hold Open Rod Fitting
CS1_CS3_7100_018
Hold Open Rod
Control Switch
Figure 17: Thrust Reverser Door Opening and Closing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-35
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
BIFURICATION LATCH SYSTEM ADJUSTMENT When engaged, the bifurcation latch uses a rod and cam mechanism to insert a locking pin through the latch. As the inner fixed structure closes, a pin on the left side is inserted into a receiver located on the right side. The pin and latch mechanism are locked together using a rod. The rod is inserted through the pin by turning, and pushing the bifurcation latch system (BLS) lever up. The bifurcation latch is locked in place when the lever is aligned with the fixed structure by pushing the lever in. The pin latch mechanism is adjustable up to 0.635 cm (0.250 in.) in either direction.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-36
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
BLS Left Hand Latch Fitting
Right Hand Latch Fitting
0.635 cm (0.250 in.) Retracted
0.635 cm (0.250 in.) Extended Nominal PIN LATCH ADJUSTABILITY
Adjustable Rod Pin Latch Receptacle
BLS Handle
Adjusting Nut
1
LOWER BLS IN OPEN POSITION
LOWER BLS IN CLOSED (LOCKED) POSITION
NOTE 1
To adjust the pin latch a distance of 0.635 cm (0.250 in.).
CS1_CS3_7100_005
Rod (engaged)
Figure 18: Thrust Reverser Door Bifurcation Latch System Adjustment Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-37
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-11 Turbine Exhaust System
78-11 TURBINE EXHAUST SYSTEM GENERAL DESCRIPTION The turbine exhaust system helps control the flow of the exhaust gases. It consists of an exhaust nozzle and centerbody. The centerbody is made up of the forward centerbody and aft centerbody. The exhaust nozzle and centerbody are bolted to the aft engine flange. Crossflow blockers, mounted on the exhaust nozzle at the 2 o’clock and 10 o'clock positions, enhance performance by reducing secondary air flow disturbances across the pylon area. Fireseal fingers, also known as turkey feather seals are mounted on the top of the exhaust nozzle to prevent flames from exiting or entering the core compartment area. During installation, two alignment pins on the forward centerbody facilitate proper clocking of the nozzle. Drain holes help remove any fluid accumulation in the turbine exhaust.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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71-38
CS130_71_L2_TES_General.fm
71 - Power Plant 78-11 Turbine Exhaust System
Exhaust Nozzle Fireseal Fingers
Aft Centerbody
Fwd Centerbody
Crossflow Blocker
Alignment Pins
CS1_CS3_7161_001
Drain Holes
Figure 19: Turbine Exhaust System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-39
CS130_71_L2_TES_General.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
71-40 ENGINE MOUNTS AND PYLON GENERAL DESCRIPTION The engine mounts attach the engine to the aircraft pylon. The nacelles do not carry any significant structural loads. Thrust loads are transferred through the engine mounts. Mounts are disconnected to remove the engine. The function of the engine mount system is to transmit engine and nacelle loads to the pylon. The pylon then transfers these loads to the wing. The main sources of engine mount loads are engine thrust, aerodynamic pressures, and inertia loads from maneuvers and gusts. Thrust links transfer thrust loads to the aft mount and prevent any bending of the engine core. All mounts are designed to be failsafe. Excess clearances in holes along secondary load paths allow the secondary paths to remain inactive, while the primary load paths are intact.
COMPONENT LOCATION Engine mounts located on the fan and turbine exhaust cases are used to attach the engine to the aircraft and pylon. The engine mount system consists of the following components: •
Forward engine mount
•
Aft engine mount
•
Thrust links
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-40
CS130_71_L2_EMP_General.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
A
B
A FWD ENGINE MOUNT
Right Thrust Link
B AFT ENGINE MOUNT
CS1_CS3_7120_001
Left Thrust Link
Figure 20: Engine Mounts and Pylon Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-41
CS130_71_L2_EMP_General.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
DETAILED COMPONENT INFORMATION FORWARD ENGINE MOUNT The forward engine mount is attached to the engine fan case, and is connected to the pylon by four bolts. Shear pins aid in aligning the mount with the pylon during engine installation. The forward engine mount supports side and vertical loads using a two-link arrangement with bolts loaded in shear. The main components of the forward mount assembly are the: •
Main beam
•
Side links
•
Shear pins
•
Failsafe bolt
The side and vertical loads at the front mount couple with the aft mount loads to support overall engine pitch and yaw. Side links provide the primary load paths from the fan case into the front beam. The front mount is attached to the pylon with four bolts and two shear pins, which transmit vertical and shear loads into the pylon. The mount bolts use captive barrel nuts to ease removal.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-42
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Fan Case
Fwd Mount Main Beam
Shear Pins
Fwd Mount Bolt Holes (4x)
Failsafe Bolt
PYLON Side Link Bolts (4x)
Fwd Mount Bolts (4x)
Captive Barrel Nuts FWD ENGINE MOUNT
CS1_CS3_7121_001
Side Link
Figure 21: Forward Engine Mount Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-43
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
AFT ENGINE MOUNT The aft engine mount is attached to the engine turbine exhaust case (TEC) and reacts to engine fore-aft, side, vertical, and roll loads. The main components of the aft mount assembly are: •
Main beam
•
Side links
•
Shear pins
•
Failsafe link
•
Balance beam
The two side links are assembled to the beam with four shear bolts. Each side link is attached to the turbine exhaust case with one shear bolt. The center link and the two attaching bolts are referred to as the failsafe link and bolts. THRUST LINKS Two thrust links attach the compressor intermediate case (CIC) to the aft mount through a balance beam. The thrust links transfer thrust loads to the aft mount. Any load differential in the two thrust links are balanced around the center pivot of the balance beam. The failure of any one link on the rear mount, including the thrust links, cause the system to transfer loads to a secondary load path.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-44
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Shear Pins Aft Mount Main Beam Thrust Links Failsafe Bolts
Balance Beam
Aft Mount Bolt Holes (4x)
Failsafe Link AFT ENGINE MOUNT
Side Links
CS1_CS3_7121_002
Aft Mount Bolts (4x)
Figure 22: Aft Engine Mount and Thrust Links Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-45
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
PRACTICAL ASPECTS ENGINE REMOVAL BOOTSTRAP EQUIPMENT The engine is removed using bootstrap equipment. The equipment consists of forward and aft bootstrap assemblies. Each assembly consists of arms and pylon mount attach points. Hoists, fitted to each arm, are used to raise and lower the engine cradle. The arms and pylon mount attach points are assembled and attached to the pylon fittings using quick-release pins.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-46
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
A
Forward Bootstrap Assembly Arms B
Quick-Release Pins
Hoist
A
B
CS1_CS3_7122_001
Aft Bootstrap Assembly Arms
Quick-Release Pins
Figure 23: Engine Removal Bootstrap Equipment Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-47
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
ENGINE CRADLE AND TRANSPORT STAND The engine cradle sits on a transport stand so that the engine can be moved. To remove the engine, the transport stand and cradle are positioned under the engine. The bootstrap hoists attach to a cradle that holds the engine. The cradle is raised up to the engine using the hoists and is attached to the engine. When the engine is disconnected from the pylon, the cradle and engine are lowered back onto the transport stand.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-48
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Hoist
Transport Stand
CS1_CS3_7122_002
Cradle
Figure 24: Engine Cradle, and Transport Stand Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-49
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
ELECTRICAL HARNESSES Electrical connectors are located at three positions on the forward pylon. There is a bracket with four connectors on each side of the engine fan case, and one bracket with three connectors on top of the engine near the precooler. All of these connectors need to be disconnected using soft jaw pliers prior to removing the engine. The generator power feeder harness connects to a terminal block, mounted near the precooler. The terminal block is accessed by removing the cover from the cooling conduit. The ring terminals need to be removed from the terminals and marked prior to removing the engine.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-50
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
D
A C
A RIGHT FAN ELECTRICAL DISCONNECTS
C VARIABLE FREQUENCY GENERATOR CABLE CONNECTS
B LEFT FAN ELECTRICAL DISCONNECTS
D VARIABLE FREQUENCY GENERATOR ELECTRICAL DISCONNECTS
CS1_CS3_7150_001
B
Figure 25: Electrical Harness Disconnects Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-51
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
HYDRAULIC AND FUEL CONNECTIONS A bracket mounted on the precooler inlet duct supports three hydraulic line quick-disconnect fittings, and two fuel line connections.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-52
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Fuel line Connections
CS1_CS3_7150_002
Hydraulic line Quick-Disconnect Fittings
Figure 26: Hydraulic and Fuel Connections Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-53
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-70 Engine Drains
71-70 ENGINE DRAINS GENERAL DESCRIPTION The engine drain system collects and discharges oil, fuel, air, and hydraulic fluid from the engine through various tubes, forming a drain mast. The drain mast terminates at the latch access panel at the 6 o’clock position on the bottom of the engine cowling. The nacelle is also designed to drain fluid. Drain paths include gaps around the latch access panel, and drain holes in the lower bifurcation fixed/access panels. Components that drain through the drain mast include: •
2.5 bleed
•
High-pressure compressor (HPC) stator vane actuator
•
Low-pressure compressor (LPC) stator vane actuator
•
Variable frequency generator (VFG)
•
Starter
•
Fuel oil manifold and integrated fuel pump and control (IFPC)
•
Oil tank scupper
•
Hydraulic pump
A map of how the drain tubes are situated in the mast assists in troubleshooting. The map is located on the inside of the latch access panel door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-54
CS130_71_L2_EngineDrains_General.fm
71 - Power Plant 71-70 Engine Drains
Fuel/Oil Manifold and IFPC
2.5 Bleed
HPC SVA
Hydraulic Pump
LPC SVA
A
B Starter C
Variable Frequency Generator
COWLING LATCH ACCESS PANEL
DRAIN TUBES MAP Oil Tank 1
DRAIN TUBES MAP DRAIN TUBES MAP 1. 999-5822-1 DR11, 2.5 BLEED 2. 999-5827-1 DR11, HPC-SVA
4
3 5
7
4. 999-5825-1 DR61, VFG 6
8
A
6. 999-5829-1 DR41, FOM-FUEL 7. 999-5826-1 DR71, OIL-TANK
3. 999-5823-1 DR01, LPC-SVA
LEGEND
5. 999-5832-1 DR91, STARTER
8. 999-5821-1 DR51, HYDRAULIC
B
DRAIN MAST
C
HPC IFPC LPC SVA
High-Pressure Compressor Integrated Fuel Pump and Control Low-Pressure Compressor Stator Vane Actuator
CS1_CS3_7171_001
2
Figure 27: Engine Drains Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-55
CS130_71_L2_EngineDrains_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
71-00 ENGINE STORAGE AND PRESERVATION GENERAL DESCRIPTION Turbine engines are designed to operate on a regular basis. Corrosion of vital engine parts can occur very quickly, especially in harsh environments such as coastal areas where the engine is exposed to salt air. Storage and preservation procedures are designed to minimize the way the environment may effect the engines. Engine covers must be installed on both engines to protect engine parts during shipping and handling, and from environmental conditions. Engine covers should always be installed if the engine is outdoors for extended periods without being operated.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-56
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
Inlet Cover
Right Rear Fan Case Cover
Hydraulic Pump Cover
Integrated Fuel Pump and Control Cover
Left Rear Fan Case Cover
Exhaust Cover
Air Turbine Starter Cover
Variable Frequency Generator Cover
CS1_CS3_7100_010
Deoiler Cover
Figure 28: Engine Storage and Preservation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-57
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
ENGINE PRESERVATION CHART
•
Check fuel for water using the ASTM D4176 procedure
The engines should be operated at least weekly. For engines that are not operated at least once a week, the following engine preservation guidelines are provided.
•
If required, drain fuel from the tanks until fuel passes the test procedure
Always refer to the aircraft maintenance publication (AMP) for preservation requirements and instructions. Method I (60 days or less) Whenever engine preservation is going to be 60 days or less the following steps need to be carried out: •
Clean and degrease the engine
•
Check for oil leaks
•
If components are removed from the engine or gearbox, ensure they are properly protected and preserved from corrosion
•
Sample the engine oil for water using ASTM D6304-00 procedure, if the sample fails then flush the engine oil system
•
Drain the engine oil system
•
Change the starter oil
•
Install tarp or PVC plastic sheets over inlet cowl, fan exhaust area and engine exhaust area
•
Make a record of the engine preservation
There is a provision to extend the engine preservation an extra 60 days if the following procedures are carried out: •
Remove the engine covers
•
Service the engine oil system
•
Run the engine at idle for five minutes once the oil temperature is above 220 °F (104.4 °C)
•
Shut down the engine
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Method II (More than 60 days) Whenever engine preservation is going to extend beyond 60 days, the following steps need to be carried out: •
Drain the engine oil system
•
Replace engine oil filter and fill the oil system
•
Start the engine and run at idle or dry motor for 3 minutes using the engine starter. It is necessary that the fan turn so that oil reaches through all of the engine oil system
•
Preserve the fuel system using preservation oil
•
Drain the engine oil
•
Change the starter oil
•
Clean and preserve the gearbox
•
Place some desiccant and some relative humidity indicators inside engine covers
•
Install tarp or PVC plastic sheets over inlet cowl, fan exhaust area and engine exhaust area. Make sure the relative humidity indicators are visible through windows in the protective covers
•
Make a record of the engine preservation
TECHNICAL TRAINING MANUAL
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71-58
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
Determine engine preservation time
More than 60 days
60 days or less
1 - 7 days
or
No action necessary
Option A
Option B
Operate the engine
Preserve for 60 days or less by (METHOD I)
Day 8
Sample the engine oil by ASTM D6304-00 procedure Preserve for 60 days or less by (METHOD I) Do not flush and drain the oil
Results OK
Results not OK
Resample the engine oil by ASTM D6304-00 procedure
Day 61
Results not OK
Do option A Run engine (option available when (METHOD I) is repeated a maximum of 1 time)
Day 121
Preserve for more than 60 days by (METHOD II) NOTE 1. This chart does not specify all of the steps in the preservation method. 2. Sampling option is restricted to one 60 day period. 3. In option B, oil system flush is not necessary (ref. METHOD I) if sample test results are OK. 4. When preserving for more than 60 days, fuel system preservation is also necessary. 5. If the engine remains in storage after 60 days (METHOD I), preserve the engine again after the initial 60 day period.
CS1_CS3_7100_011
Return engine to service
Figure 29: Engine Preservation Chart Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-59
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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71-60
CS130_71_ILB.fm
ATA 72 - Engine
BD-500-1A10 BD-500-1A11
72 - Engine Fuel and Control
Table of Contents 72-00 PW1500G Engine.........................................................72-2 General Description ..........................................................72-2 Component Location ........................................................72-4 72-11 Fan Assembly...............................................................72-6 General Description ..........................................................72-6 Detailed Component Information ......................................72-8 Fan Hub .......................................................................72-8 Fan Blades...................................................................72-8 Inlet Cone.....................................................................72-8 Fan Case Assembly...................................................72-10 72-15 Fan Drive Gear System ..............................................72-12 General Description ........................................................72-12 Detailed Component Information ....................................72-14 Fan Drive Gearbox.....................................................72-14 No. 1 and No. 1.5 Bearing Support Assembly ...........72-16 72-22 Fan Intermediate Case ...............................................72-18 General Description ........................................................72-18 Detailed Component Information ....................................72-20 Fan Exit Fairing..........................................................72-20 Fan Exit Stator ...........................................................72-20 Fan Intermediate Case ..............................................72-20 No.2 Bearing and Support Assembly.........................72-20 Fan Drive Gearbox Coupling .....................................72-20 Low-Pressure Compressor Variable Stator Vanes ...72-20 72-31 Low-Pressure Compressor.........................................72-22 General Description ........................................................72-22 Detailed Component Information ....................................72-24 Low-Pressure Compressor Case...............................72-24 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Low-Pressure Compressor Stators ........................... 72-24 Integral Bladed Rotor ................................................ 72-24 No.2 Bearing ............................................................. 72-24 Tie Rods.................................................................... 72-24 72-34 Compressor Intermediate Case ................................. 72-26 General Description ....................................................... 72-26 Detailed Component Information.................................... 72-28 Compressor Intermediate Case ................................ 72-28 No. 3 Bearing Compartment ..................................... 72-28 Gearbox Drive Bevel Gear ........................................ 72-28 Fire Shield ................................................................. 72-28 Low-Pressure Compressor Exit Stator ...................... 72-28 High-Pressure Compressor Variable Inlet Guide Vanes ....................................... 72-28 72-35 High-Pressure Compressor ....................................... 72-30 General Description ....................................................... 72-30 Detailed Component Information ................................... 72-32 Rotors........................................................................ 72-32 High-Pressure Compressor Rotor Shaft ................... 72-32 Stators....................................................................... 72-32 72-41 Diffuser, Combustor and Turbine Nozzle Assembly .................................................... 72-34 General Description ....................................................... 72-34 Detailed Component Information ................................... 72-36 Diffuser Case Assembly............................................ 72-36 Combustion Chamber ............................................... 72-36 High-Pressure Turbine Nozzle Assembly ................. 72-36 72-51 High-Pressure Turbine............................................... 72-38 General Description ....................................................... 72-38
TECHNICAL TRAINING MANUAL
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72-i
CS130_M5_72_EngineTOC.fm
72 - Engine Fuel and Control
Detailed Component Information ....................................72-40 High-Pressure Turbine...............................................72-40 72-52 Turbine Intermediate Case .........................................72-42 General Description ........................................................72-42 Detailed Component Information ....................................72-44 Turbine Intermediate Case ........................................72-44 72-53 Low-Pressure Turbine ................................................72-46 General Description ........................................................72-46 Detailed Component Information ....................................72-48 Low-Pressure Turbine................................................72-48 72-54 Turbine Exhaust Case ................................................72-50 General Description ........................................................72-50 Detailed Component Information ....................................72-52 Turbine Exhaust Case ...............................................72-52 72-60 Main and Angle Gearboxes ........................................72-54 General Description ........................................................72-54 Main Gearbox ............................................................72-54 Main Gearbox Mounted Components ........................72-56 Angle Gearbox ...........................................................72-58 72-00 Borescope Access......................................................72-60 General Description ........................................................72-60 Engine Left Side Borescope Access..........................72-60 Engine Right Side Borescope Access .......................72-62 Borescope Access - Main Gearbox ...........................72-64 Borescope Access - Angle Gearbox ..........................72-64
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TECHNICAL TRAINING MANUAL
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72-ii
CS130_M5_72_EngineTOC.fm
72 - Engine Fuel and Control
List of Figures Figure 1: PW1500G Engine..................................................72-3 Figure 2: PW1500G Engine Assembly Component Location .............................................72-5 Figure 3: Fan Assembly........................................................72-7 Figure 4: Fan Hub, Blades, and Inlet Cone ..........................72-9 Figure 5: Fan Case Assembly ............................................72-11 Figure 6: Fan Drive Gear System.......................................72-13 Figure 7: Fan Drive Gearbox ..............................................72-15 Figure 8: No. 1 and No. 1.5 Bearing Support Assembly and Cross-Section ..............................72-17 Figure 9: Fan Intermediate Case........................................72-19 Figure 10: Fan Intermediate Case Cross-Section ...............72-21 Figure 11: Low-Pressure Compressor..................................72-23 Figure 12: Low-Pressure Compressor Cross-Section .........72-25 Figure 13: Compressor Intermediate Case ..........................72-27 Figure 14: Compressor Intermediate Case Cross-Section......................................................72-29 Figure 15: High-Pressure Compressor.................................72-31 Figure 16: High-Pressure Compressor Cross-Section .........72-33 Figure 17: Diffuser, Combustor, and Turbine Nozzle Assembly....................................72-35 Figure 18: Diffuser and Combustor Cross-Section...............72-37 Figure 19: High-Pressure Turbine .......................................72-39 Figure 20: High-Pressure Turbine Cross-Section.................72-41 Figure 21: Turbine Intermediate Case.................................72-43 Figure 22: Turbine Intermediate Case Cross-Section ..........72-45 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 23: Low-Pressure Turbine ........................................ 72-47 Figure 24: Low-Pressure Turbine Cross-Section................. 72-49 Figure 25: Turbine Exhaust Case ....................................... 72-51 Figure 26: Turbine Exhaust Case Cross-Section................. 72-53 Figure 27: Main Gearbox ..................................................... 72-55 Figure 28: Main Gearbox Mounted Components................. 72-57 Figure 29: Angle Gearbox.................................................... 72-59 Figure 30: Engine Left Side Borescope Access................... 72-61 Figure 31: Engine Right Side Borescope Access ................ 72-63 Figure 32: Borescope Access Gearbox and Angle Gearbox .............................. 72-65
TECHNICAL TRAINING MANUAL
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72-iii
CS130_M5_72_EngineLOF.fm
ENGINE - CHAPTER BREAKDOWN ENGINE CHAPTER BREAKDOWN
Engine
1
72 - Engine 72-00 PW1500G Engine
72-00 PW1500G ENGINE GENERAL DESCRIPTION The PW1500G series engines are high bypass ratio turbofans, utilizing two spools with an integral primary and fan duct exhaust systems. The N1 rotor consists of a single-stage low-pressure ratio fan, driven at a reduced speed through a fan drive gear system (FDGS), a 3-stage low-pressure compressor (LPC), and a 3-stage low-pressure turbine (LPT) on a common shaft. The FDGS is a planetary star gear reduction unit that takes torque from the LPT and reduces the fan speed to 30% of the LPT speed. The LPC has variable inlet guide vanes, which are positioned automatically by the electronic engine control (EEC) to optimize engine operation. The N2 rotor consists of the high-pressure compressor (HPC) with eight compressor stages, and a high-pressure turbine (HPT) consisting of two stages. The HPC has variable inlet guide vanes and variable stator vanes in the 1st, 2nd, and 3rd stages which are positioned automatically by the EEC to optimize engine operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-2
CS130_72_L2_PW1500G_General.fm
72 - Engine 72-00 PW1500G Engine
3 Stage Low-Pressure Compressor
8 Stage High-Pressure Compressor
2 Stage High-Pressure Turbine
3 Stage Low-Pressure Turbine
Fan Rotor
2
2.5
3
4
4.5 5
1
LEGEND NOTE
Fan Rotor/FDGS N1 Rotor N2 Rotor
1
Engine station numbers.
CS1_CS3_7200_009
4.9
Figure 1: PW1500G Engine Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-3
CS130_72_L2_PW1500G_General.fm
72 - Engine 72-00 PW1500G Engine
COMPONENT LOCATION The PW1500G engine is made up of the following modules: •
Fan rotor
•
Fan case
•
Fan drive gear system (FDGS)
•
Fan intermediate case (FIC)
•
Low-pressure compressor (LPC)
•
Compressor intermediate case (CIC)
•
High-pressure compressor (HPC)
•
Diffuser and combustor
•
High-pressure turbine (HPT)
•
Turbine intermediate case (TIC)
•
Low-pressure turbine (LPT)
•
Turbine exhaust case (TEC)
•
Angle gearbox
•
Main gearbox (MGB)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-4
CS130_72_L2_PW1500G_ComponentLocation.fm
72 - Engine 72-00 PW1500G Engine
Fan Case
FDGS
CIC DIFF/COMB FDGS FIC HPC HPT LPC LPT TEC TIC
Compressor Intermediate Case Diffuser/Combustor Fan Drive Gear System Fan Intermediate Case High-Pressure Compressor High-Pressure Turbine Low-Pressure Compressor Low-Pressure Turbine Turbine Exhaust Case Turbine Intermediate Case
Angle Gearbox
FAN ROTOR
FIC
LPC
CIC
Main Gearbox
HPC
DIFF/ COMB
ENGINE MODULES
HPT
TIC
LPT
TEC
CS1_CS3_7200_010
LEGEND
Figure 2: PW1500G Engine Assembly Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-5
CS130_72_L2_PW1500G_ComponentLocation.fm
72 - Engine 72-11 Fan Assembly
72-11 FAN ASSEMBLY GENERAL DESCRIPTION The fan assembly is composed of the fan rotor and the fan case assembly. The fan rotor is located inside the fan case assembly. It turns clockwise and produces 90% of the engine total thrust. The fan case assembly directs the fan air through and around the engine core.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-6
CS130_72_L2_FanAssembly_General.fm
72 - Engine 72-11 Fan Assembly
Fan Case
CS1_CS3_7200_008
Fan Rotor
Figure 3: Fan Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-7
CS130_72_L2_FanAssembly_General.fm
72 - Engine 72-11 Fan Assembly
DETAILED COMPONENT INFORMATION FAN HUB The fan hub retains the fan blades on its circumference. The hub has attachment points on the front face for the inlet cone and fan rotor trim balance weights. The no. 1 fan blade slot is located between two no.1 marks on the hub. Numbering of the fan blades and fairings is in the direction of rotation, beginning at the no. 1 blade position. Details on fan rotor balancing can be found in ATA 77, Engine Indicating. FAN BLADES The fan rotor has 18 wide-chord hollow aluminum fan blades. The hollow aluminum fan blades have titanium leading edges for wear resistance and foreign object debris (FOD) tolerance. A protective polyurethane black paint coating provides erosion protection for the fan blade airfoil. Each fan blade has a dovetail shaped root that fits into slots on the fan hub. Composite wear strips are bonded to the fan blade dovetail sides. Composite fan blade fairings are located between the fan blades to provide an inner flow path surface. Each fan blade fairing is held in place by a fan blade retaining pin, inserted into pin lugs on the fan hub. A spacer is installed between each blade and the hub to minimize wear, and provide a tight fit. The fan blades are held in place on the hub by a front locking ring. INLET CONE The composite inlet cone is installed in front of the fan hub to smooth the airflow into the engine. The inlet cone cover provides access to the inlet cone retaining hardware. The inlet cone is removed to gain access to the fan blades lock ring.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-8
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-11 Fan Assembly
Titanium Leading Edge
Aluminum Fan Blade
Fan Hub Fan Blade Fairing
Composite Wear Strip
Fan Blade Fairing Pin
Fan Blade Spacer
Lock Ring
1
Fan Blade Fairing Pin Lugs
Inlet Cone 1
Fan Rotor Trim Balance Weights Attachment Flange
Inlet Cone Attachment Flange
CS1_CS3_7211_002
No. 1 Fan Blade Marks
Inlet Cone Cover
Figure 4: Fan Hub, Blades, and Inlet Cone Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-9
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-11 Fan Assembly
FAN CASE ASSEMBLY
Precooler Duct Inlet
The fan case serves as the structural link between the inlet cowl and the engine core by providing an attachment point to the airframe. It consists of a one-piece composite case with a titanium mount ring attached around the aft outer surface. A Kevlar containment ring surrounds the fan case.
The precooler duct inlet is a four-piece titanium casting mounted at the rear top of the fan case. It is used to direct fan air into the pylon mounted precooler, provides seal lands for the thrust reverser (T/R) door firewall seal, and sealing of the upper nacelle flow path. The precooler duct inlet is attached to the titanium mount ring at the rear of the case. The mount ring provides the front mount attachment points for the engine mounts, and includes a V-groove around the entire circumference, providing alignment and support for the thrust reverser doors.
Fan case components include: •
Fan blade rub strips
•
Acoustic panels
•
Ice liners
•
Precooler duct inlet
•
Fan exit guide vanes (FEGV)
•
Fan exit liner segments
Fan Exit Guide Vanes On the inner surface of the fan case, 44 hollow aluminum fan exit guide vanes (FEGVs) extend rearward to the fan intermediate case (FIC). The FEGVs straighten the fan air and provide structural support between the FIC and the fan case assembly. A polyurethane coating covers most of the FEGV surfaces to protect against corrosion and erosion.
Fan Blade Rub Strips
Fan Exit Liner Segments
The fan blade rub strips provide an abradable surface for the fan blade to rub against when the blades elongate under power. They also protect the fan blades from contacting the fan case and ensure a tight clearance between the case and the rub strips. The surface of the rub strips can be refurbished if any significant rubbing occurs.
Six fan exit liner segments, immediately aft of the FEGVs, surround the low-pressure compressor (LPC). Stationary louvers in the fan exit liner segments provide an exit path for the 2.5 bleed air into the fan air stream.
Acoustic Panels
The segments can be removed individually for component access and inspection of lines and fittings.
Acoustic panels, forward of the fan rub strip, and acoustic liner segments aft of the fan blades, decrease noise generated from the fan. Ice Liners Replaceable ice liners aft of the fan, provide protection from ice for the case structure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-10
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-11 Fan Assembly
Fan Case
Engine Mount
Precooler Duct Inlet A
Fan Exit Liner Segments
A FORWARD ENGINE MOUNT V-Groove
Fan Blade Rub Strip Acoustic Panels
Ice Liner Rear Acoustic Liner Segment
Titanium Mount Ring
B FAN EXIT GUIDE VANE
CS1_CS3_7221_001
B
Figure 5: Fan Case Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-11
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
72-15 FAN DRIVE GEAR SYSTEM GENERAL DESCRIPTION The fan drive gear system (FDGS) is a reduction gear assembly that uses torque from the low-pressure turbine (LPT) N1 to turn the fan assembly at approximately one third the speed of the LPT. This results in a lower fan speed and a higher low-pressure compressor (LPC) speed, thus increasing compressor efficiency. The FDGS is located between the fan rotor assembly and the fan intermediate case (FIC). It is radially and axially supported by the no.1 and no. 1.5 bearing support assembly. The LPC shaft connects to the input of the FDGS. The input speed is reduced by the fan drive gearbox (FDG) and then transmitted to a gearbox shaft connected to the fan rotor hub.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-12
CS130_72_L2_FDGS_General.fm
72 - Engine 72-15 Fan Drive Gear System
Support Assemblies Fan Drive Gear System B
Input from Low-Pressure Turbine Shaft A
Fan Drive Gearbox
A
No. 1 and No. 1.5 Bearing Support
B FAN DRIVE GEAR SYSTEM
CS1_CS3_7210_001
Gearbox Shaft
Figure 6: Fan Drive Gear System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-13
CS130_72_L2_FDGS_General.fm
72 - Engine 72-15 Fan Drive Gear System
DETAILED COMPONENT INFORMATION FAN DRIVE GEARBOX The fan drive gearbox (FDG) is the most aft section of the FDGS. It houses the sun gear, the five star gears, and the rotating ring gear halves. The FDG takes the torque from the low-pressure turbine (LPT) FDG titanium coupling to turn a sun gear. This sun gear turns five stationary star gears against an outer ring gear set. The outer ring gear set is turned by the star gears and is connected to the gearbox shaft, which connects to the fan hub. Each star rotates around its own journal bearing mounted in the carrier. The journal bearings provide the rotational hinge for the star gears. The oil manifold supplies oil to no. 1 and no. 1.5 bearing support assembly, which supplies the lubrication for the FDG gears and bearings. The torque frame links the carrier to the flex support that is bolted to the fan intermediate case (FIC). This flex support ensures the FDGS remains aligned with the FDG coupling. The FDG coupling connects the LPC shaft to the FDGS. The FDG coupling is a specially designed flex shaft that is used to reduce pressure on the FDGS. It transmits the torque from the LPC to the FDGS input.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-14
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
Ring Gear Set Halves
Gearbox Shaft
Journal Bearings
Sun Gear
Torque Frame
FDG Coupling Gearbox Shaft No. 1 and No. 1.5 Bearing Support Assembly
Star Gears
Carrier
Oil Manifold
Flex Support
CS1_CS3_7210_003
Fan Drive Gearbox
Figure 7: Fan Drive Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-15
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
NO. 1 AND NO. 1.5 BEARING SUPPORT ASSEMBLY The no. 1 and no. 1.5 bearing support assembly supports the gearbox shaft using two tapered roller bearings. The no. 1 bearing is oil damped to absorb fan vibration. The aft section of the bearing support assembly is attached to the fan intermediate case (FIC).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-16
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
No. 1 Bearing Support No. 1 and No. 1.5 Bearing Support
Gearbox Shaft
No. 1.5 Bearing CS1_CS3_7215_001
No. 1 Bearing
Figure 8: No. 1 and No. 1.5 Bearing Support Assembly and Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-17
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-22 Fan Intermediate Case
72-22 FAN INTERMEDIATE CASE GENERAL DESCRIPTION The fan intermediate case (FIC) provides the support structure between the fan case, the fan drive gear system (FDGS), and the low-pressure compressor (LPC). The fan drive gearbox (FDG) is housed inside the FIC. It also provides the aerodynamic flow path between the fan rotor and the LPC. The fan exit fairing splits the air between the fan duct and the low-pressure compressor. Air entering the fan exit stators is directed to the LPC variable stator vanes.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-18
CS130_72_L2_FIC_General.fm
72 - Engine 72-22 Fan Intermediate Case
Fan Intermediate Case
Fan Exit Stator
CS1_CS3_7222_011
Fan Exit Fairing
Figure 9: Fan Intermediate Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-19
CS130_72_L2_FIC_General.fm
72 - Engine 72-22 Fan Intermediate Case
DETAILED COMPONENT INFORMATION
LOW-PRESSURE COMPRESSOR VARIABLE STATOR VANES
FAN EXIT FAIRING
The low-pressure compressor (LPC) variable stator vanes are located at the FIC aft end and mate with the LPC case front end. The LPC variable stators vane improve engine stability and operation.
The fan exit fairing is installed on the front flange of the fan intermediate case (FIC). It splits the airflow between the engine core section, and the fan. FAN EXIT STATOR The fan exit stator is installed on the front flange of the FIC, just inside of the splitter. It straightens the fan air flow and directs it to the low-pressure compressor (LPC) variable vane stator. FAN INTERMEDIATE CASE The fan intermediate case (FIC) keeps the fan drive gear system (FDGS) and the LPC in alignment, and also provides support for the no. 1 and no. 1.5 bearing support assembly. NO.2 BEARING AND SUPPORT ASSEMBLY The no. 2 bearing and support assembly is attached to the rear center of the FIC. It absorbs rotor radial movement, and is oil damped on the outer race to decrease vibration. FAN DRIVE GEARBOX COUPLING The fan drive gearbox coupling (FDG coupling) is located in the heart of the FIC. Its aft end sits inside of bearing no. 2 and extends forward to couple with the fan drive gearbox system (FDGS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-20
CS130_72_L3_FIC_DetailedComponentInfo.fm
72 - Engine 72-22 Fan Intermediate Case
Fan Exit Fairing
Fan Intermediate Case
Fan Exit Stator
Low-Pressure Compressor Variable Stator Vane
Low-Pressure Compressor
No. 1 and No. 1.5 Bearing Support
No. 2 Bearing and Assembly Support
Low-Pressure Compressor Rotor Hub
CS1_CS3_7222_001
Fan Drive Gearbox Coupling
Figure 10: Fan Intermediate Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-21
CS130_72_L3_FIC_DetailedComponentInfo.fm
72 - Engine 72-31 Low-Pressure Compressor
72-31 LOW-PRESSURE COMPRESSOR GENERAL DESCRIPTION The low-pressure compressor (LPC) has three stages. Each stage has an integral bladed rotor (IBR). An IBR consists of a disc and blades manufactured as a single unit, resulting in increased efficiency and less weight than conventional bladed discs. An annular bleed valve at the rear of the LPC compressor case, designated as the 2.5 bleed valve, provides operational stability for the LPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-22
CS130_72_L2_LPC_General.fm
72 - Engine 72-31 Low-Pressure Compressor
Low-Pressure Compressor Case
2.5 Bleed Valve
CS1_CS3_7231_001
Stage 1 Integral Bladed Rotor (IBR)
Figure 11: Low-Pressure Compressor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-23
CS130_72_L2_LPC_General.fm
72 - Engine 72-31 Low-Pressure Compressor
DETAILED COMPONENT INFORMATION LOW-PRESSURE COMPRESSOR CASE The 2.5 bleed valve is mounted on the LPC case. Removable fan exit liner segments cover the LPC case. LOW-PRESSURE COMPRESSOR STATORS There are two stages of stator vanes in the LPC. Each stage consists of fixed aluminum segment stators. INTEGRAL BLADED ROTOR The three integral bladed rotors (IBRs) are tied to the LPC rotor hub outer diameter using tie rods. The LPC rotor hub is splined onto the low-pressure turbine (LPT) shaft, just aft of the oil damped no. 2 bearing. It also couples to the fan drive gearbox (FDG) coupling. NO.2 BEARING The no. 2 bearing supports the LPC rotor hub, the FDG coupling, and the LPT shaft. TIE RODS There are 48 tie rods that connect the IBRs to the LPC rotor hub. They are located on the second IBR.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-24
CS130_72_L3_LPC_DetailedComponentInfo.fm
72 - Engine 72-31 Low-Pressure Compressor
Fan Exit Liner Segment
Low-Pressure Compressor Case
Stator Vanes
2.5 Bleed Valve
Integral Bladed Rotors
Tie Rod
No. 2 Bearing
Low-Pressure Compressor Rotor Hub
Low-Pressure Turbine Shaft
CS1_CS3_7231_002
Fan Drive Gearbox Coupling
Figure 12: Low-Pressure Compressor Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-25
CS130_72_L3_LPC_DetailedComponentInfo.fm
72 - Engine 72-34 Compressor Intermediate Case
72-34 COMPRESSOR INTERMEDIATE CASE GENERAL DESCRIPTION The compressor intermediate case (CIC) acts as the support structure for the low-pressure compressor (LPC), and the high-pressure compressor (HPC). The CIC carries the thrust loads from the engine through the two thrust links attachment points to the aft engine mounts. The two thrust link attachment points are on the CIC at the 2 o’clock and 10 o’clock positions. The CIC also houses the no. 3 bearing compartment. The case structure includes nine struts and an opening for a gearbox drive shaft that transmits torque to the angle gearbox. The case is equipped with an opening for the discharge of the 2.5 bleed air. The hollow struts provide support for the no.3 bearing compartment. The struts interior provide access to the no.3 bearing for lubrication supply and scavenge lines. The struts are numbered clockwise beginning at the 1 o’clock position. A table lists all the struts internal passage usage.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-26
CS130_72_L2_CIC_General.fm
72 - Engine 72-34 Compressor Intermediate Case
Strut 9
Strut 1
Strut 8
High-Pressure Compressor Variable Inlet Guide Vanes
Strut 2
Strut 7
Strut 3
Strut 6
Strut 4
Thrust Link Attachment Point
Strut 5 VIEW LOOKING FORWARD
COMPRESSOR INTERMEDIATE CASE STRUTS
Low-Pressure Compressor Exit Stator
No. 3 Bearing Compatment
2.5 Bleed Air Openings
Used For
1
No. 2 bearing buffer air supply
2
Breather tube
3
No. 3 bearing oil supply
4
N1 probe
5
No. 3 bearing oil scavenge
6
Gearbox drive shaft
7
Empty
8
No. 3 bearing buffer air supply
9
No. 3 bearing buffer air supply
CS1_CS3_7234_001
Strut No.
Figure 13: Compressor Intermediate Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-27
CS130_72_L2_CIC_General.fm
72 - Engine 72-34 Compressor Intermediate Case
DETAILED COMPONENT INFORMATION
HIGH-PRESSURE COMPRESSOR VARIABLE INLET GUIDE VANES
COMPRESSOR INTERMEDIATE CASE
The high-pressure compressor variable inlet guide vanes are located at the aft end of the CIC. They are adjusted during engine operation to optimize engine performance.
The compressor intermediate case (CIC) includes the no. 3 bearing, the 2.5 bleed air outlet, and the thrust reverser (T/R) door inner fixed structure (IFS) support. NO. 3 BEARING COMPARTMENT The no. 3 bearing compartment is located in the aft section of the compressor intermediate case (CIC). The no. 3 bearing supports the front of the high-pressure compressor’s (HPC) rotor shaft axial movement and absorbs vibrations. GEARBOX DRIVE BEVEL GEAR The gearbox drive bevel gear is located at the 8 o’clock position in the no. 3 bearing compartment. The no. 3 bearing bevel gear engages the gearbox drive bevel gear to supply torque to the angle gearbox drive shaft. This shaft drives the angle gearbox, which in turn, supplies torque to the main gearbox. FIRE SHIELD The fire shield is attached to the outer perimeter of the CICs aft section. It is a metal containment shield used to protect the LPC structure from fire originating in the engine core compartment. The outer section of the shield acts as a support for the thrust reverser (T/R) door inner fixed structure (IFS). LOW-PRESSURE COMPRESSOR EXIT STATOR The low-pressure compressor (LPC) exit stator segments are located at the front of the CIC. The LPC exit stators divert air from the LPC, and direct it towards the HPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-28
CS130_72_L3_CIC_DetailedComponentInfo.fm
72 - Engine 72-34 Compressor Intermediate Case
Gearbox Drive Bevel Gear
No. 3 Bearing Bevel Gear
No. 3 Bearing Compartment
High-Pressure Compressor Rotor Shaft Angle Gearbox Drive Shaft Gearbox Drive Bevel Gear Support
High-Pressure Compressor High-Pressure Compressor Variable Inlet Guide Vanes
Low-Pressure Compressor No. 2.5 Exit Stators Bleed Air Outlet
Fire Shield
Angle Gearbox Thrust Reverser Door Inner Fixed Structure Support
CS1_CS3_7234_003
Low-Pressure Compressor
Figure 14: Compressor Intermediate Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-29
CS130_72_L3_CIC_DetailedComponentInfo.fm
72 - Engine 72-35 High-Pressure Compressor
72-35 HIGH-PRESSURE COMPRESSOR GENERAL DESCRIPTION The high-pressure compressor (HPC) increases the speed and pressure of the primary gaspath air. The HPC is driven by the high-pressure turbine (HPT) through the HPC rotor shaft. The first seven rotor stages are integral bladed rotors (IBRs). The first three stator stages are variable and actuated through unison rings on the outside of the HPC case. The forward HPC case is split horizontally and mated with a one piece aft ring case. The HPC supplies 4th and 8th stage air to the bleed air system. The 4th stage bleed air port is located on the HPC. Engine cooling is supplied by 4th stage air through cooling air ports on the HPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-30
CS130_72_L2_HPC_General.fm
72 - Engine 72-35 High-Pressure Compressor
4th Stage Bleed Air Port
High-Pressure Compressor Rotor Shaft
Stator Vane Unison Rings
Aft Ring Case
Split Front Case Stage 1 Integral Bladed Rotor
CS1 CS3 7235 001 CS1_CS3_7235_001
4th Stage Cooling Air Ports
Figure 15: High-Pressure Compressor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-31
CS130_72_L2_HPC_General.fm
72 - Engine 72-35 High-Pressure Compressor
DETAILED COMPONENT INFORMATION ROTORS There are eight rotor stages in the high-pressure compressor (HPC). The first seven rotor stages are integral bladed rotors (IBRs). The 8th stage rotor is a standard blade and disk assembly. HIGH-PRESSURE COMPRESSOR ROTOR SHAFT The rotor stage hubs are held together on the HPC rotor shaft. The HPC front hub acts as the forward stop for the IBRs and the HPC rear hub secures the 8th stage rotor. This rotor hub is held by the HPC rotor shaft nut on the HPC rotor shaft. STATORS The first three stator stages of the HPC are variable, while the remaining four stages are fixed stages.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-32
CS130_72_L3_HPC_DetailedComponentInfo.fm
72 - Engine 72-35 High-Pressure Compressor
HPC Front Hub
Integral Bladed Rotors
Variable Stator Vanes
8th Stage Rotor
Fixed Stator Vanes
HPC Rear Hub
LEGEND HPC
HPC Rotor Shaft
HPC Rotor Shaft Nut
CS1_CS3_7235_002
Compressor Intermediate Case
High-Pressure Compressor
Figure 16: High-Pressure Compressor Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-33
CS130_72_L3_HPC_DetailedComponentInfo.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
72-41 DIFFUSER, COMBUSTOR AND TURBINE NOZZLE ASSEMBLY GENERAL DESCRIPTION The diffuser case, combustion chamber, and turbine nozzle assemblies work together to provide the internal combustion that drives the turbines. Compressed air flows from the high-pressure compressor (HPC) exit stator into the diffuser and then into the combustor section. The airflow is straightened and diffused, reducing the velocity, while maintaining pressure before it enters the combustion chamber. The diffuser case has the bleed air ports for the 6th stage bleed air, cooling air, and cowl anti-ice (CAI). The combustion chamber inner perimeter and the turbine nozzles are attached to the diffuser inner case. The combustion chamber consists of an inner and outer chamber liner assembly. The diffuser case also has provisions for fuel nozzles and igniter plugs. Metered fuel is supplied to the combustion chamber through 16 fuel nozzles, where it is mixed with the air. The heat of the combustion process causes the airflow to expand and accelerate rearward. Turbine nozzle guide vanes then direct the high-temperature, high-velocity gases out of the combustion chamber to drive the turbines. The diffuser outer case supports the main gearbox (MGB) using two mount brackets, located at the 3 o'clock and 9 o'clock position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-34
CS130_72_L2_DCTNA_General.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
Turbine Nozzle Assembly
Inner Liner
Outer Liner COMBUSTION CHAMBER
Diffuser Outer Case High-Pressure Compressor Exit Stator Gearbox Mount Bracket (2x)
Diffuser Inner Case DIFFUSER CASE ASSEMBLY
6th Stage Cooling Air Port
CS1_CS3_7240_001
Fuel Nozzle (16x)
Figure 17: Diffuser, Combustor, and Turbine Nozzle Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-35
CS130_72_L2_DCTNA_General.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
DETAILED COMPONENT INFORMATION DIFFUSER CASE ASSEMBLY The diffuser inner case attaches to a common diffuser outer case flange with the high-pressure compressor (HPC). The inner case attaches to the high-pressure turbine (HPT) flange. Two igniter plugs located at the 4 o’clock and 5 o'clock position extend into the combustion chamber. COMBUSTION CHAMBER The HPC air enters the diffuser and flows through the hood and cooling holes around the inner and outer liners of the combustion chamber. A bulkhead is used to attach both combustion chamber inner and outer liners, and the hood. Fuel nozzles inject fuel through orifices in the bulkhead where swirlers spin the air/fuel mixture inside the combustion chamber. HIGH-PRESSURE TURBINE NOZZLE ASSEMBLY The turbine nozzle assembly attaches to the cooling air duct, which is bolted with the combustion chamber inner liner flange, to the inner diffuser case. The nozzle assembly has 32 air-cooled thermal-barrier coated guide vanes between the inner and outer combustion chamber liners. Located at the rear of the combustion chamber, they direct the hot gases to the 1st stage high-pressure turbine blades. Each hollow vane is cooled by diffuser air and use cooling air exit holes on the airfoil external surface to reduce vane temperature.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-36
CS130_72_L3_DCTNA_DetailedComponentInfo.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
Fuel Nozzle
High-Pressure Compressor
Igniter Plug
Diffuser Case High-Pressure Compressor Exit Stator
High-Pressure Turbine
Outer Liner
Hood Combustion Chamber
1st Stage Vane Bulkhead Swirler
Inner Liner
CS1_CS3_7241_001
Inner Diffuser Case
Cooling Air Holes
Figure 18: Diffuser and Combustor Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-37
CS130_72_L3_DCTNA_DetailedComponentInfo.fm
72 - Engine 72-51 High-Pressure Turbine
72-51 HIGH-PRESSURE TURBINE GENERAL DESCRIPTION The high-pressure turbine (HPT) provides the rotational force to drive the HPC by extracting energy from the hot combustion gases. The HPT case provides containment of the HPT rotors. Cooling air flows through the hollow 2nd stage stator vanes to keep them cool.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-38
CS130_72_L2_HPT_General.fm
72 - Engine 72-51 High-Pressure Turbine
CS1_CS3_7251_001
2nd Stage Vane Cooling Air
Figure 19: High-Pressure Turbine Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-39
CS130_72_L2_HPT_General.fm
72 - Engine 72-51 High-Pressure Turbine
DETAILED COMPONENT INFORMATION HIGH-PRESSURE TURBINE Each of the two turbine rotor stages contain 44 blades coated with a thermal barrier and internally cooled. The 1st stage hub engages a spline in the HPC rear hub at the front. The 2nd stage hub is attached by a nut on the HPC rotor shaft at the rear. The blades are attached to their hubs in fir tree type slots and held in position with retaining plates. The blades outer air seal segments are ground as a set with the turbine blades to maximize rotor tip sealing. Internally cooled 2nd stage stator vanes are mounted inside the case between the two rotor stages. Fan air from the active clearance control (ACC) system cools the HPT case, thus decreasing its inner diameter and increasing the turbine efficiency. The HPT case has mounting points for the turbine ACC manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-40
CS130_72_L3_HPT_DetailedComponentInfo.fm
72 - Engine 72-51 High-Pressure Turbine
HPT Case
Turbine Active Clearance Control (ACC) Manifold Mount
HPT 2nd Stage Vane
Turbine Rotor Stages
HPC Rotor Shaft
HPT 1st Stage Hub
HPT 2nd Stage Hub
CS1_CS3_7251_002
HPC Rear Hub
Figure 20: High-Pressure Turbine Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-41
CS130_72_L3_HPT_DetailedComponentInfo.fm
72 - Engine 72-52 Turbine Intermediate Case
72-52 TURBINE INTERMEDIATE CASE GENERAL DESCRIPTION The turbine intermediate case (TIC) is used to turn the high-pressure turbine (HPT) airflow to align it with the counter rotating low-pressure turbine (LPT). The no. 4 roller bearing housing assembly is located within the TIC. The turbine intermediate case houses a turbine stator assembly which connects to the intermediate case using seven support rods. The turbine stator assembly is composed of 14 airfoils that direct HPT airflow to the LPT rotor.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-42
CS130_72_L2_TIC_General.fm
72 - Engine 72-52 Turbine Intermediate Case
Support Rod (7x)
Turbine Stator Assembly Airfoils
CS1_CS3_7251_003
No. 4 Roller Bearing Housing Assembly
Figure 21: Turbine Intermediate Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-43
CS130_72_L2_TIC_General.fm
72 - Engine 72-52 Turbine Intermediate Case
DETAILED COMPONENT INFORMATION TURBINE INTERMEDIATE CASE The turbine intermediate outer case holds the turbine intermediate inner case in a center position with seven rods. These rods run through the turbine stator vanes and provides support to the no.4 roller bearing housing assembly. It is oil damped in order to absorb rotor vibration, and supports the HPT drive shaft axially and radially. Heat shields are provided as thermal barriers to protect the no. 4 roller bearing housing assembly. Buffer air from the 4th stage provides sealing and cooling for the no. 4 bearing compartment, and prevents coking of engine oil. Lubrication and air cooling tubes are routed through the hollow turbine stator vanes to and from the no.4 roller bearing housing assembly. An outer heat shield provides a thermal barrier to the TICs outer case.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-44
CS130_72_L3_TIC_DetailedComponentInfo.fm
72 - Engine 72-52 Turbine Intermediate Case
Turbine Intermediate Case Rod (7x)
High-Pressure Turbine
Outer Heat Shield
Turbine Stator Vanes (14x)
Heat Shield Heat Shield Turbine Intermediate Inner Case
Heat Shield
No. 4 Roller Bearing Housing Assembly
High-Pressure Turbine Drive Shaft
CS1_CS3_7252_001
Low-Pressure Turbine Drive Shaft
Figure 22: Turbine Intermediate Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-45
CS130_72_L3_TIC_DetailedComponentInfo.fm
72 - Engine 72-53 Low-Pressure Turbine
72-53 LOW-PRESSURE TURBINE GENERAL DESCRIPTION The low-pressure turbine (LPT) drives the LPT shaft. The LPT case provides an attachment for the active clearance control (ACC) manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-46
CS130_72_L2_LPT_General.fm
72 - Engine 72-53 Low-Pressure Turbine
Active Clearance Control Manifold
Low-Pressure Turbine Case
CS1_CS3_7253_001
Low-Pressure Turbine Shaft
Figure 23: Low-Pressure Turbine Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-47
CS130_72_L2_LPT_General.fm
72 - Engine 72-53 Low-Pressure Turbine
DETAILED COMPONENT INFORMATION LOW-PRESSURE TURBINE The low-pressure turbine (LPT) consists of 3 rotor stages and 2 stages of stator vanes. The LPT drive shaft runs through the center of the engine. The LPT 2nd stage rotor hub is splined to the LPT drive shaft. The low-pressure turbine (LPT) 1st stage and 3rd stage rotor disks are bolted together on the LPT 2nd stage hub using tie rods. The LPT 1st stage rotor blades are hollow, while the 2nd and 3rd stage rotor blades are solid alloy. The LPT 2nd and 3rd stator vanes are segment arrangement type. Fan air is sprayed around the exterior of the LPT case from the active clearance control (ACC) cooling manifold. This reduces the LPT case diameter, reducing the turbine blades tip clearance, thus improving fuel efficiency. Insulation segments, located on the inside periphery of the LPT, protect the LPT case from heat.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-48
CS130_72_L3_LPT_DetailedComponentInfo.fm
72 - Engine 72-53 Low-Pressure Turbine
Insulation Segments
Low-Pressure Turbine Case
Active Clearance Control Cooling Manifold
Low-Pressure Turbine 1st Stage Rotor Blade
Low-Pressure Turbine 2nd Stage Stator Vane Low-Pressure Turbine 3rd Stage Rotor Blade
Low-Pressure Turbine 3rd Stage Stator Vane
Low-Pressure Turbine 2nd Stage Rotor Blade Tie Rod
Low-Pressure Turbine Drive Shaft
Low-Pressure Turbine 2nd Stage Rotor Hub
Low-Pressure Turbine 3rd Stage Rotor Disk
CS1_CS3_7253_002
Low-Pressure Turbine 1st Stage Rotor Disk
Figure 24: Low-Pressure Turbine Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-49
CS130_72_L3_LPT_DetailedComponentInfo.fm
72 - Engine 72-54 Turbine Exhaust Case
72-54 TURBINE EXHAUST CASE GENERAL DESCRIPTION The turbine exhaust case (TEC) is the aft most engine structural element. It is formed by the outer duct surface, the inner duct surface, eleven airfoil shaped struts, and the no.5 and no.6 bearing housings. The TEC is attached to the rear flange of the low-pressure turbine (LPT) at the front end, and to the exhaust system at the aft end. It forms a transition duct between the LPT and the exhaust case. Two ground handling bosses are provided at the 3 o’clock and 6 o'clock position. The TEC also houses the aft engine mount lugs at its top surface. Eight exhaust gas temperature (EGT) probes bosses are provided on the case outer surface. Oil supply and scavenge tubes use the hollow interior of two lower struts to reach no. 5 and no. 6 bearing housings.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-50
CS130_72_L2_TEC_General.fm
72 - Engine 72-54 Turbine Exhaust Case
Aft Engine Mount Lugs
Strut (11x)
Inner Duct Surface
Exhaust Gas Temperature Probe (8x)
No. 6 Bearing Housing Outer Duct Surface
No. 5 and No. 6 Bearing Oil Supply Tube
Ground Handling Boss (2x) No. 5 and No. 6 Bearing Oil Scavenge Tube
CS1_CS3_7254_001
No. 5 Bearing Housing
Figure 25: Turbine Exhaust Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-51
CS130_72_L2_TEC_General.fm
72 - Engine 72-54 Turbine Exhaust Case
DETAILED COMPONENT INFORMATION TURBINE EXHAUST CASE The turbine exhaust case (TEC) and struts attach to the no. 5 and no. 6 bearing housings. Heat shields are installed to protect both bearing housings from excessive hot exhaust gas flow. The no. 5 and no. 6 bearings are roller bearings that provide the radial support for the LPT shaft. Both bearings are oil-damped to absorb vibration. The no. 5 bearing rests against its forward seat and is properly kept separated from no.6 bearing by a shaft spacer. The no. 6 bearing retaining nut is installed on the low-pressure turbine (LPT) shaft to keep both bearing assemblies together. The three aft engine mount lugs carry part of the engine weight and loads to the pylon structure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-52
CS130_72_L3_TEC_DetailedComponentInfo.fm
72 - Engine 72-54 Turbine Exhaust Case
Aft Engine Mount Lugs
Low-Pressure Turbine
Turbine Exhaust Case
Strut
Heat Shield
No. 6 Bearing Housing No. 6 Bearing
No. 6 Bearing Retaining Nut
No. 5 Bearing Housing Low-Pressure Turbine Shaft No. 5 Bearing Seat
No. 5 Bearing
Shaft Spacer
CS1_CS3_7254_002
Heat Shield
Figure 26: Turbine Exhaust Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-53
CS130_72_L3_TEC_DetailedComponentInfo.fm
72 - Engine 72-60 Main and Angle Gearboxes
72-60 MAIN AND ANGLE GEARBOXES The angle gearbox mechanically connects the high-pressure compressor (N2) and the main gearbox through the angle gearbox layshaft.
GENERAL DESCRIPTION MAIN GEARBOX The main gearbox (MGB) housing is a cast aluminum structure mounted to the engine core. Accessories are mounted on the forward and aft faces of the gearbox. The gearbox housing has the following components: •
Replaceable carbon seals
•
De-oiler
•
N2 crank pad port
•
Core drain plugs
•
Permanent magnet alternator generator (PMAG) drive
The MGB uses gears and shafts to drive accessories mounted to it. The angle gearbox layshaft drive is the rotational input to the MGB. The N2 speed is reduced before reaching the angle gearbox in order to have the MGB input drive in the proper rotational speed range. The gearbox aft face is axially supported by a rod link assembly that connects to the turbine intermediate case (TIC). A centering puck is installed on the gearbox top surface to help align the gearbox to the engine when it is installed. Two mounting links are used to mount the gearbox to the diffuser case. A heat shield protects the gearbox from engine core radiating heat. It is clipped at the front side and bolted to brackets at the aft end top surface of the gearbox.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-54
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
Heat Shield Left Mounting Link
Centering Puck
Air Turbine Starter Carbon Seal
Self Contained Deoiler N2 Cranking Port Pad Right Mounting Link Rod Link Assembly
Integrated Fuel Pump and Control Carbon Seal
Angle Gearbox Layshaft Drive Hydraulic Pump Carbon Seal
Deoiler Carbon Seal Permanent Magnet Alternator Generator (PMAG) Drive
Variable Frequency Generator Carbon Seal
CS1_CS3_7261_002
Core Drain Plugs
Lubrication and Scavenge Oil Pump Drive Pad
Figure 27: Main Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-55
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
MAIN GEARBOX MOUNTED COMPONENTS The following components are mounted on the main gearbox: •
Hydraulic pump
•
Lubrication and scavenge oil pump (LSOP)
•
Oil control module (OCM)
•
Variable frequency generator (VFG)
•
Permanent magnet alternator generator (PMAG)
•
Air turbine starter
•
Fuel oil manifold
•
Integrated fuel pump and control (IFPC)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-56
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
Fuel/Oil Manifold
Integrated Fuel Pump and Control
Angle Gearbox Main Gear Box
Air Turbine Starter
Hydraulic Pump Lubrication and Scavenge Oil Pump
Oil Control Module Variable Frequency Generator
CS1_CS3_7261_001
Permanent Magnet Alternator Generator
Figure 28: Main Gearbox Mounted Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-57
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
ANGLE GEARBOX The angle gearbox mechanically connects the HPC (N2) and the main gearbox (MGB) through the layshaft. The angle gearbox is installed at the 8 o'clock position on the compressor intermediate case (CIC). The shaft is covered by two layshaft covers. The forward layshaft cover front end is secured to the angle gearbox, and the rear cover aft end is secured to the MGB. Both layshaft covers slide one over the other at the mating center area to allow for thermal growth during engine operation. When the engine is being started, the air starter drives the MGB, which is connected to the angle gearbox. The angle gearbox drives the HPC. When the engine is running, the HPC drives the MGB through the angle gearbox.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-58
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
Angle Gearbox
Layshaft
CS1_CS3_7260_001
Layshaft Covers
Figure 29: Angle Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-59
CS130_72_L2_MAG_General.fm
72 - Engine 72-00 Borescope Access
72-00 BORESCOPE ACCESS GENERAL DESCRIPTION
ENGINE LEFT SIDE BORESCOPE ACCESS
The borescope permits visual inspection of internal gaspath parts without the need for disassembly.
Table 1: Left Side Borescope Access Ports and Inspection Areas
A borescope probe is inserted in one of several access ports to inspect parts for damage, cracks, wear and missing material.
PORT
LOCATION
AP-LPC 3
9:30 position
LPC 2nd stage stator vanes and 3rd stage IBR blades.
AP-HPT 1
9 o’clock position
HPT 1st stage rotor blades, 2nd stage vanes, and 2nd stage rotor blades.
AP-LPT 1
8:30 position
LPT 1st stage rotor blades, LPT 1st stage vanes, and LPT 2nd stage rotor blades.
AP-LPT 2
8:30 position
LPT 2nd stage rotor blades, LPT 2nd stage vanes, and LPT 3rd stage rotor blades.
AP-DIFF 1
11 o’clock position
Combustion chamber inner and outer liners, bulkhead, fuel nozzles, HPT 1st stage rotor blades, and HPT 1st stage vanes.
Some damage to blades or stators can be blended in-situ using borescope blade blending. Consult the Aircraft Maintenance Publication (AMP) for airfoil damage and blend limitations. The following tables show the borescope access ports and inspection areas.
INSPECTION AREA
CS1_CS3_TB72_003
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-60
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
LPC 3
DIFF 1
HPT 1 LPT 1
CS1_CS3_7200_005
LPT 2
Figure 30: Engine Left Side Borescope Access Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-61
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
ENGINE RIGHT SIDE BORESCOPE ACCESS Table 2: Right Side Borescope Access Ports and Inspection Areas PORT
LOCATION
INSPECTION AREA
AP-LPC 1
3:30 position
LPC variable inlet stator vanes 1st stage integral bladed rotor (IBR) airfoils.
AP-LPC 2
4 o’clock position
LPC 1st stage stator vanes and 2nd stage IBR airfoils.
AP-HPC 1
4 o’clock position
HPC variable inlet guide vanes and 1st stage IBR airfoils.
AP-HPC 2
4 o’clock position
HPC 1st stage variable stator vanes (VSV) and 2nd stage IBR airfoils.
AP-HPC 3
3:30 position
HPC 2nd stage VSV and 3rd stage IBR airfoils.
AP-HPC 4
4 o’clock position
HPC 3rd stage VSV and 4th stage IBR airfoils.
AP-HPC 5
4.o’clock position
HPC4th stage stator and 5th stage IBR airfoils.
AP-HPC 6
4 o’clock position
HPC 5th stage stator and 6th stage IBR airfoils.
AP-HPC 7
2 o’clock position
HPC 6th stage stator and 7th stage IBR airfoils.
AP-HPC 8
2 o’clock position
HPC 7th stage stator and 8th stage rotor blade.
Igniter Plug Ports
3:30 and 4 o’clock positions
Combustion chamber inner and outer liner, bulkhead, fuel nozzles, HPT 1st stage rotor blades and vanes.
AP-TIC 1
1:30 position
HPT 2nd stage rotor blades. LPT 1st stage rotor blades. CS1_CS3_TB72_004
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-62
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
TIC 1
HPC 8 HPC 7
HPC 6
Igniter Upper
HPC 2 HPC 3 HPC 4 HPC 1
Igniter Lower
LPC 2
LPC 1
CS1_CS3_7200_006
HPC 5
Figure 31: Engine Right Side Borescope Access Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-63
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
BORESCOPE ACCESS - MAIN GEARBOX The gearbox also has two access ports to facilitate inspection. An N2 crank pad is used to turn the N2 rotor during the borescope procedure. BORESCOPE ACCESS - ANGLE GEARBOX The angle gearbox has one borescope access port for internal inspection.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-64
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
N2 Crank Pad Cover
Borescope Port
Borescope Port
Borescope Port Angle Gearbox
CS1_CS3_7200_007
Main Gearbox
Figure 32: Borescope Access - Gearbox and Angle Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-65
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-66
CS130_72_ILB .fm
ATA 73 - Engine Fuel and Control
BD-500-1A10 BD-500-1A11
73 - Engine Fuel and Control
Table of Contents 73-11 Distribution....................................................................73-2 General Description ..........................................................73-2 Component Location ........................................................73-4 Component Information ....................................................73-6 Integrated Fuel Pump and Control...............................73-6 Fuel Filter .....................................................................73-8 Fuel Nozzles ..............................................................73-10 Detailed Description .......................................................73-12 Motive Flow Fuel........................................................73-12 Servo Fuel..................................................................73-12 Fuel Filter ...................................................................73-12 Metered Fuel..............................................................73-12 Fuel/Oil Heat Exchanger Operation ...........................73-12 73-20 Control System ...........................................................73-14 General Description ........................................................73-14 Electronic Engine Control ..........................................73-16 Fuel Control Description ............................................73-18 Prognostics and Health Management Unit ................73-20 Component Location ......................................................73-22 P2/T2 Probe...............................................................73-22 Burner Pressure Sensor ............................................73-22 Electronic Engine Control Unit ...................................73-22 Prognostics and Health Management Unit ................73-22 Data Storage Unit ......................................................73-22 Permanent Magnet Alternator Generator...................73-22 P2.5/T2.5 Probe.........................................................73-24 T3 Probe ....................................................................73-24 Component Information ..................................................73-26 Electronic Engine Control ..........................................73-26 Data Storage Unit ......................................................73-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Prognostics and Health Management unit ................ 73-26 Permanent Magnet Alternator Generator ................. 73-28 Detailed Component Information ................................... 73-30 Electronic Engine Control Power .............................. 73-30 Electronic Engine Control Fault Detection ................ 73-32 Prognostics and Health Management Unit................ 73-34 Data Storage Unit...................................................... 73-36 Controls and Indications ................................................ 73-38 Throttle Quadrant Assembly ..................................... 73-38 Engine Panel............................................................. 73-38 Engine Fire Panel...................................................... 73-38 Indications ................................................................. 73-40 Detailed Description ...................................................... 73-42 Aircraft and Electronic Engine Control Interface ....... 73-42 Engine Sensor Interface............................................ 73-44 Engine Operating Limits............................................ 73-46 Overspeed Protection, Shaft Shear Protection and Thrust Control Malfunction ................................. 73-48 Integrated Fuel Pump and Control Operation - Start ........................................... 73-50 Integrated Fuel Pump and Control Operation - Run ............................................ 73-52 Normal Engine Shutdown ........................................ 73-54 Emergency Shutdown ............................................... 73-56 Engine Power Management...................................... 73-58 Engine Power Settings and Modes ........................... 73-60 Automatic Power Reserve......................................... 73-62 N1 Synchronization ................................................... 73-64 Fault Reporting and Testing Interface....................... 73-66 Monitoring and Tests ...................................................... 73-68 CAS Messages ......................................................... 73-69
TECHNICAL TRAINING MANUAL
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73-i
CS130_M5_73_EngFuelCtlTOC.fm
73 - Engine Fuel and Control
Practical Aspects.............................................................73-72 Onboard Maintenance System ..................................73-72 73-30 Indicating ....................................................................73-76 General Description ........................................................73-76 Component Location ......................................................73-78 Controls and Indications .................................................73-80 Detailed Description .......................................................73-82 Monitoring and Tests.......................................................73-84 CAS Messages ..........................................................73-85
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-ii
CS130_M5_73_EngFuelCtlTOC.fm
73 - Engine Fuel and Control
List of Figures Figure 1: Fuel Distribution System General Description.......73-3 Figure 2: Fuel Distribution System Component Location .....73-5 Figure 3: Integrated Fuel Pump and Control ........................73-7 Figure 4: Fuel Filter ..............................................................73-9 Figure 5: Fuel Nozzles........................................................73-11 Figure 6: Fuel Distribution System Detailed Description ....73-13 Figure 7: Engine Control System General Description.......73-15 Figure 8: Electronic Engine Control....................................73-17 Figure 9: Fuel Control.........................................................73-19 Figure 10: Prognostics and Health Management Unit..........73-21 Figure 11: Engine Control System Component Location .....73-23 Figure 12: P2.5/T2.5 Probe and T3 Probe Location.............73-25 Figure 13: Electronic Engine Control, Data Storage Unit, and Prognostics and Health Management Unit..........73-27 Figure 14: Permanent Magnet Alternator Generator ............73-29 Figure 15: Electronic Engine Control Power.........................73-31 Figure 16: Electronic Engine Control Fault Detection...........73-33 Figure 17: Prognostics and Health Management Unit..........73-35 Figure 18: Data Storage Unit................................................73-37 Figure 19: Fuel Control System Controls ............................73-39 Figure 20: Fuel Control System Indications.........................73-41 Figure 21: Aircraft and Electronic Engine Control Interface..................................................73-43 Figure 22: Engine Sensor Interface......................................73-45 Figure 23: Electronic Engine Control Engine Operating Limits ..................................................73-47 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 24: Overspeed, Shaft Shear Protection, and Thrust Control Malfunction........................... 73-49 Figure 25: Integrated Fuel Pump and Control Operation - Engine Start ........................ 73-51 Figure 26: Integrated Fuel Pump and Control Operation - Engine Run ......................... 73-53 Figure 27: Normal Engine Shutdown ................................... 73-55 Figure 28: Emergency Shutdown........................................ 73-57 Figure 29: Engine Bleed Management and Idle Speed ....... 73-59 Figure 30: Engine Power Settings and Modes..................... 73-61 Figure 31: Automatic Power Reserve .................................. 73-63 Figure 32: N1 Synchronization............................................. 73-65 Figure 33: Electronic Engine Control Fault Reporting and Testing Interface ......................... 73-67 Figure 34: OMS EEC System Test ...................................... 73-73 Figure 35: Power Assurance Test........................................ 73-75 Figure 36: Fuel Control Indicating System........................... 73-77 Figure 37: Fuel Control Indicating System Component Location........................................... 73-79 Figure 38: Fuel Control Indications ...................................... 73-81 Figure 39: Fuel Control Indicating Detailed Description....... 73-83
TECHNICAL TRAINING MANUAL
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73-iii
CS130_M5_73_EngFuelCtlLOF.fm
ENGINE FUEL AND CONTROL - CHAPTER BREAKDOWN ENGINE FUEL AND CONTROL CHAPTER BREAKDOWN
Distribution
1
Control System
2
Indicating
3
73 - Engine Fuel and Control 73-11 Distribution
73-11 DISTRIBUTION GENERAL DESCRIPTION The engine fuel distribution system supplies fuel to the engine for combustion and for air system actuator operation. The aircraft fuel system supplies fuel to the integrated fuel pump and control (IFPC). A fuel oil manifold provides mounting for the IFPC, fuel flow transmitter and the fuel filter. The manifold simplifies maintenance by minimizing the number of external fuel lines needed to connect the components. The IFPC has integral pumps driven by the engine gearbox: •
Centrifugal, low-pressure boost pump
•
High-pressure main gear pump to provide pressurized fuel
•
Motive flow gear pump
The filtered fuel then flows to the main pump where it is pressurized. Some of the pressurized fuel is supplied to the air system actuators for engine cooling and stability control. The main pump also supplies the fuel to the fuel metering valve (FMV). The FMV is controlled by the electronic engine control (EEC) to provide the correct fuel for a given thrust setting. The metered fuel flow is measured by the fuel flow meter. The fuel is then sent to the flow divider. The flow divider valve (FDV) allows fuel to flow to the primary fuel nozzles for engine start. As the engine accelerates to idle, the flow divider valve provides flow to the secondary fuel nozzles.
Fuel from the aircraft is supplied to the IFPC boost pump under pressure. Should the aircraft fuel system pumps fail, the boost pump is capable of suctioning fuel from the aircraft fuel tanks. The boost pump provides fuel at low-pressure to the motive flow pump and to a fuel/oil heat exchanger (FOHX). The motive flow pump provides fuel to operate the aircraft fuel system ejector pumps. The FOHX heats engine fuel to prevent icing and cools the engine oil for proper engine bearing seal operation. The heated fuel then passes through the fuel filter, which is installed in a housing attached to the fuel manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-2
CS130_73_L2_Distribution_General.fm
73 - Engine Fuel and Control 73-11 Distribution
INTEGRATED FUEL PUMP AND CONTROL PRESSURE-REGULATING VALVE Fuel From Main Tank
Boost Pump MAIN PUMP
LEGEND
ACC EEC FDV FMV FOHX HPC LPC TM
Engine Oil Fuel Metered Fuel Motive Flow Pressurized Fuel Pump Discharge Servo Fuel Servo Fuel Return Active Clearance Control Electronic Engine Control Flow Divider Valve Fuel Metering Valve Fuel/Oil Heat Exchanger High-Pressure Compressor Low-Pressure Compressor Torque Motor
FMV
MOTIVE FLOW PUMP
FDV
FUEL/OIL MANIFOLD To Primary Manifold To Secondary Manifold
To Motive Flow System
To Engine Oil
FOHX
TURBINE HPC LPC ACC STATOR STATOR AIR VANES VANE VALVE FUEL FILTER
2.5 BLEED VALVE
EEC
CS1_CS3_7311_001
From Engine Oil
FUEL FLOW METER
Figure 1: Fuel Distribution System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-3
CS130_73_L2_Distribution_General.fm
73 - Engine Fuel and Control 73-11 Distribution
COMPONENT LOCATION
Fuel/Oil Manifold
The engine fuel distribution system consists of the following components:
The fuel/oil manifold is located on the front face of the gearbox on the left side of the engine.
•
Integrated fuel pump and control
•
Fuel/oil heat exchanger
•
Fuel/oil heat exchanger bypass valve
•
Fuel filter
•
Fuel nozzles
•
Fuel supply manifolds
•
Fuel/oil manifold
•
Fuel flow meter
Fuel Flow Meter The fuel flow meter is attached to the fuel oil manifold on the front face of the main gearbox at the 9 o’clock position.
Integrated Fuel Pump and Control The integrated fuel pump and control (IFPC) is mounted on the fuel/oil manifold on the front face of the main gearbox at the 8 o’clock position. Fuel/Oil Heat Exchanger and Fuel/Oil Heat Exchanger Bypass Valve The fuel/oil heat exchanger (FOHX) and fuel/oil heat exchanger bypass valve are located at the 12 o’clock position on the compressor intermediate case. Fuel Filter The fuel filter is located below the IFPC on the fuel/oil manifold. Fuel Nozzles and Primary and Secondary Fuel Manifolds The fuel nozzles and primary and secondary fuel manifolds are located on the circumference of the combustor.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-4
CS130_73_L2_Distribution_ComponentLocation.fm
73 - Engine Fuel and Control 73-11 Distribution
Fuel/Oil Heat Exchanger Bypass Valve (FOHXBV)
Fuel/Oil Heat Exchanger (FOHX) Fuel Supply Manifolds
Fuel Flow Meter
Fuel Nozzles
Fuel/Oil Manifold
Fuel Filter
CS1_CS3_7311_002
Integrated Fuel Pump and Control (IFPC)
Figure 2: Fuel Distribution System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-5
CS130_73_L2_Distribution_ComponentLocation.fm
73 - Engine Fuel and Control 73-11 Distribution
COMPONENT INFORMATION INTEGRATED FUEL PUMP AND CONTROL The IFPC supplies metered fuel based on inputs from the EEC in response to changes in thrust demand and environmental conditions. The IFPC performs the following major functions: •
Supply motive fuel for aircraft fuel system
•
Meter fuel for combustion
•
Pressurize fuel supply
•
Fuel shutoff for engine shutdown
•
Control primary and secondary flow distribution to fuel nozzles
•
Supply pressurized servo fuel for air system actuators
•
Bypass excess fuel flow
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-6
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
MOTIVE FLOW
METERED FUEL FLOW
FUEL PRESSURE REGULATION
FUEL SHUTOFF
PRIMARY AND SECONDARY MANIFOLD PRESSURIZATION
INTEGRATED FUEL PUMP AND CONTROL
SERVO FUEL
CS1_CS3_7311_003
EXCESS FUEL BYPASS
Figure 3: Integrated Fuel Pump and Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-7
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
FUEL FILTER The main fuel filter element is used to remove contaminants from the fuel sent from the IFPC. The canister type disposable fuel filter is installed in a housing on the fuel oil manifold below the IFPC. The fuel filter bypass valve opens when the filter clogs and the differential pressure exceeds 25 psid. The filter housing is equipped with a drain plug for draining the housing prior to removing the filter.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-8
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
Filter Housing Drain
Fuel Filter Bypass Valve
CS1_CS3_7311_004
Fuel Filter Housing
Figure 4: Fuel Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-9
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
FUEL NOZZLES There are 16 fuel nozzles, 10 duplex, and 6 simplex. They are arranged in a pattern of 4 duplex, 4 simplex, 6 duplex, and 2 simplex nozzles. The IFPC and fuel/oil manifold send fuel to the primary and secondary fuel nozzle manifolds that supply the fuel nozzles. The duplex nozzles contain both primary fuel nozzles and secondary fuel nozzles. The simplex nozzles only contain secondary fuel nozzles. All of the fuel nozzles contain check valves to prevent fuel in the primary and secondary fuel nozzle manifolds from draining into the engine following an engine shutdown. These valves in the primary and secondary circuits of the fuel nozzles shut when the fuel system pressure drops below a minimum threshold during shutdown conditions, preventing fuel from draining into the combustor. During an engine start, the IFPC delivers fuel to the primary manifold to provide combustion for initial spool up of the engine. During engine acceleration to ground idle, the FDV in the IFPC distributes fuel flow to both the primary and secondary nozzles. An alignment pin on each nozzle prevents incorrect installation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-10
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
Primary Flow Secondary Flow
Simplex Nozzle
Secondary Manifold
Duplex Nozzle Alignment Pin
Primary Manifold
DUPLEX NOZZLE
Secondary Flow FUEL NOZZLES (AFT LOOKING FORWARD)
FUEL NOZZLES ON COMBUSTION CHAMBER
SIMPLEX NOZZLE
CS1_CS3_7311_005
Alignment Pin
Figure 5: Fuel Nozzles Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-11
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
DETAILED DESCRIPTION MOTIVE FLOW FUEL Motive flow for the aircraft fuel system is provided by a motive flow gear pump downstream of the centrifugal boost pump exit. The inlet of the motive flow pump is protected from fuel contamination by a screen with a bypass valve. The outlet of the motive flow pump is protected from overpressure by a high-pressure relief valve. SERVO FUEL The servo pressure regulator adjusts the pressure to a suitable level for actuator operation. Servo fuel is supplied to the turbine active clearance control (ACC) valve, high-pressure compressor (HPC) stator vane actuator, low-pressure compressor (LPC) stator vane actuator, and the 2.5 bleed valve actuator. The actuators are bi-directional fully modulated linear actuators used to optimize overall engine performance and to prevent engine compressor surges. The servo fuel is then returned from the actuators upstream of the fuel filter before being recirculated in the IFPC.
The metered fuel output of the FMV then flows to the fuel flow meter as long as the windmill bypass valve (WBV) remains closed. The fuel flow meter sends a flow signal to the EEC for fuel flow indication. Fuel is supplied to the minimum pressure-shutoff valve (MPSOV) and flow divider valve (FDV) for distribution to the fuel nozzles. When the engine is shutdown, the WBV opens and connects fuel from upstream of the minimum pressure-shutoff valve (MPSOV) with the fuel filter inlet. The MPSOV closes and fuel is shut off from the engine. FUEL/OIL HEAT EXCHANGER OPERATION The fuel/oil heat exchanger maintains the fuel temperature above 0°C (32°F) at the fuel filter inlet during steady-state operation by transferring engine oil heat to the fuel. The EEC controls the fuel/oil heat exchanger bypass valve (FOHXBV) based on the fuel temperature sensor signal. As the temperature decreases, the EEC opens the FOHXBV to supply more hot oil to warm the fuel. If the fuel side of the heat exchanger is blocked, a bypass valve opens to divert fuel around the heat exchanger.
FUEL FILTER The fuel filter differential pressure sensor measures the pressure drop across the fuel filter. If the fuel filter is completely clogged, a bypass valve opens to divert fuel around the fuel filter. The EEC monitors the fuel filter and provides an L(R) ENG FUEL FILTER advisory message when the fuel filter clogs. An accompanying INFO message indicates a filter clogging or an actual bypass condition. METERED FUEL The pressure-regulating valve (PRV) modulates to provide a constant fuel flow to the FMV regardless of the FMV position. The EEC computes the required metered fuel using the fuel temperature sensor for compensation. In the case of a fuel temperature sensor failure, the EEC uses a default fuel temperature of 75°C (167°F). Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-12
CS130_73_L3_Distribution_DetailedDescription.fm
73 - Engine Fuel and Control 73-11 Distribution
INTEGRATED FUEL PUMP AND CONTROL LEGEND
FTS HPC LPC MPSOV T/M WBV
¨3
Fuel Flow Differential Pressure Sensor
PRESSURE REGULATING VALVE Boost Pump WBV
MAIN PUMP FTS FMV MOTIVE FLOW PUMP
Servo Pressure Regulator MPSOV
FDV
FUEL/OIL MANIFOLD
To Primary Manifold To Secondary Manifold
Screen To Motive Flow
FUEL FLOW METER T/M
¨3
FOHX
TURBINE HPC LPC ACC STATOR STATOR AIR VANES VANE VALVE
From Engine Oil FOHXBV To Engine Oil
FUEL FILTER
2.5 BLEED VALVE
EEC
CS1_CS3_7311_007
ACC EEC FDV FMV FOHX FOHXBV
Engine Oil Fuel From Fuel Main Tank Metered Fuel Motive Flow Pressurized Fuel Pump Discharge Servo Fuel Servo Fuel Return Active Clearance Control Electronic Engine Control Flow Divider Valve Fuel Metering Valve Fuel/Oil Heat Exchanger Fuel/Oil Heat Exchanger Bypass Valve Fuel Temperature Sensor High-Pressure Compressor Low-Pressure Compressor Minimum Pressure-Shutoff Valve Torque Motor Windmill Bypass Valve Bypass Valve
Figure 6: Fuel Distribution System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-13
CS130_73_L3_Distribution_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
73-20 CONTROL SYSTEM GENERAL DESCRIPTION The engine is controlled by a full authority digital engine control (FADEC). A dual-channel electronic engine control (EEC) controls the engine thrust, using N1 as a reference, in accordance with the position of the throttle quadrant assembly (TQA) thrust lever, environmental conditions, and the amount of bleed air being extracted from the engine. This is done by the EEC using the data received from both aircraft systems and engine sensors, together with the engine trim data inputs from the data storage unit (DSU).
A gearbox-driven permanent magnet alternator generator (PMAG) provides electrical power to the EEC during normal operation. The EEC channel A and channel B, each have their own processor, power supply, program memory, and inputs and outputs. Either EEC channel is capable of controlling all of the critical engine functions: •
Engine fuel control
•
Engine starting control
The FADEC consists of three main components:
•
Engine stability
•
Electronic engine control (EEC)
•
Engine limit control
•
Prognostics and health management unit (PHMU)
•
Data storage unit (DSU)
During normal operation, one channel is in control and the other channel is in standby. Both channels calculate and provide the same commands, however, only the outputs from the channel in control are active. In case of an output failure on the channel in control, the equivalent output from the standby channel is used and controlled by the standby channel.
The data storage unit (DSU) is installed on the EEC and used for storage of engine configuration and identification data, including information related to the engine’s specific serial number. The prognostics and health management unit (PHMU) is a single channel line replaceable unit (LRU) that monitors engine parameters to provide the engine health status, vibration, auxiliary oil pressure sensing, and oil debris monitoring. The software in the EEC provides the logic required for engine control, fault detection and accommodation. The engine indication and crew alerting system (EICAS) display parameters, and engine diagnostics and maintenance data to the onboard maintenance system (OMS). The EEC also provides health monitoring information from the PHMU to the health management unit (HMU). All of the FADEC components have fieldloadable software.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-14
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
Aircraft Systems
DMC 1 and DMC 2
EICAS SYN OMS
THROTTLE QUADRANT ASSEMBLY
HMU ELECTRONIC ENGINE CONTROL
DSU
Channel A Engine Fuel Control Engine Starting Control Engine Stability Engine Limit Protection Channel B
ENGINE SENSORS DMC DSU
Data Concentrator Unit Module Cabinet Data Storage Unit ARINC 429 CAN BUS
PERMANENT MAGNET ALTERNATOR GENERATOR
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
NOTE Left engine shown. Right engine similar.
CS1_CS3_7321_012
LEGEND
Figure 7: Engine Control System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-15
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
ELECTRONIC ENGINE CONTROL
Cowl Anti-Ice
The EEC channels are normally powered by the PMAG when engines are running but are powered by ESSENTIAL DC BUSES when the engines and the PMAG fails or the engine is shutdown. The left EEC channel A is powered by DC ESS BUS 3 and channel B is powered by DC ESS BUS 1. The right EEC channel A is powered by DC ESS BUS 3, and channel B is powered by DC ESS BUS 2.
The cowl anti-ice (CAI) valves are controlled by the EEC to provide anti-icing for the inlet cowl. They are also opened to improve HPC stability during engine start and at low engine power settings.
Fuel Control The EEC schedules engine thrust in response to thrust lever angle from the throttle quadrant assembly (TQA), aircraft inputs from the integrated air management system (IAMS), flight management system (FMS) and engine sensor inputs. Engine N1 and N2 limits are protected at all times and the exhaust gas temperature (EGT) is limited at start. The main input for fuel flow scheduling is thrust lever angle from the TQA.
Air Systems The engine interfaces with the IAMS to provide the correct bleed demand based on air conditioning and wing anti-ice demands. Thrust Reverser The EEC controls the deploy and stow functions of the thrust reverser actuation system (TRAS). The EEC control is based on thrust lever position from the TQA when the aircraft is on the ground.
Additional parameters the EEC uses to calculate fuel flow are ambient pressure (Pambient), inlet air temperature (T2), and pressure (P2), burner pressure (Pb), exhaust gas temperature (EGT), N1 speed and N2 speed. Starting and Ignition System The control system provides automatic starting sequencing of the starter air valve for the air turbine starter, igniters, and hydraulic pump depressurization during the starting phase. Engine Airflow Control System The EEC controls bleed valves, and the low-pressure compressor (LPC) and high-pressure compressor (HPC) stator vane system for engine stability and operability. The EEC also controls the turbine active clearance control systems for improved engine efficiency.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-16
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
THROTTLE QUADRANT ASSEMBLY
ELECTRONIC ENGINE CONTROL DC ESS BUS 3
FUEL CONTROL DSU STARTING AND IGNITION SYSTEM
Channel A ENGINE AIRFLOW CONTROL
PERMANENT MAGNET ALTERNATOR GENERATOR DC ESS BUS 1
COWL ANTI ICE Channel B AIR SYSTEMS
THRUST REVERSER ENGINE SENSORS DSU
Data Storage Unit ARINC 429 CAN BUS
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
NOTE Left engine shown. Right engine similar.
CS1_CS3_7321_032
LEGEND
Figure 8: Electronic Engine Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-17
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
FUEL CONTROL DESCRIPTION Each engine is controlled by a thrust lever on the TQA. The thrust lever controls both forward and reverse thrust. The ENG switches provide engine run and shutdown commands for normal operation. The ENGINE panel has a START switch to initiate the start sequence in manual mode.
The fuel scheduling and limit protection is based on the following sensors: •
N1 speed
•
N2 speed
•
Ambient air pressure (Pamb)
The EEC uses information from the air data system for engine control. The air data is validated against data received from the engine sensors P2T2 and Pamb. Engine sensor data is used in the event the air data is not available, however, engine performance is degraded.
•
Inlet air temperature (T2)
•
Inlet pressure (P2)
•
Burner pressure (Pb)
The flight management system (FMS) provides thrust requirements based on the flight plan information entered. The FMS also provides an interface to manually control the thrust reference.
•
Exhaust gas temperature (EGT)
If a fire is detected, the engine is automatically shut down when the L(R) FIRE PBA is pressed.
The integrated air management system supplies the bleed air requirements from the aircraft systems. The thrust requirement is adjusted to ensure that the minimum bleed air requirements are always met. When the thrust requirements are set, the autothrottle can be engaged to control the thrust levers automatically, based on the FMS flight plan inputs. The EEC provides fuel scheduling and engine start commands to the integrated fuel pump and control (IFPC). The IFPC provides metered fuel for combustion and servo fuel for control of the engine air systems.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Engine rotor speeds are calculated by the EEC using the N1 and N2 signals. The EEC schedules fuel according to the calculated N1 target and limits fuel if N1 or N2 limits are reached. If an overspeed is detected on N1 or N2, the engine is shutdown. Burner pressure (Pb) is sent to the EEC to schedule fuel, detect stalls, and hung starts. The exhaust gas temperature (EGT) is monitored by the EEC. During engine start, the maximum EGT is limited. The EGT is used for temperature exceedance determination, light off detection, hot start abort and engine health monitoring. The EEC limits fuel flow when redline limits are reached for N1, N2 or burner pressure. In addition to redline limiting, if the high or low rotor speed exceeds the overspeed limit of 105% N1 or 105% N2, the EEC will command an overspeed solenoid in the integrated fuel pump and control unit (IFPC) to cut engine fuel flow.
TECHNICAL TRAINING MANUAL
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73-18
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
L ENG
F IR E BTL 1
BTL 2
AVAIL
AVAIL
FMS AIR DATA
L ENGINE FIRE PANEL AUTO THROTTLE ENGINE CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
DMC 1 and DMC 2
INTEGRATED AIR MANAGEMENT SYSTEM
ENGINE PANEL
MAX
ELECTRONIC ENGINE CONTROL IDLE
INTEGRATED FUEL PUMP AND CONTROL
IDLE
L ENG
ON
FF FF OFF
FF FF OFF
Servo Fuel
N1 N2 Pamb T2 P2 Pb EGT
R ENG
ON
Metered Fuel For Combustion
THROTTLE QUADRANT ASSEMBLY LEGEND DMC
NOTE
Data Concentrator Unit Module Cabinet ARINC 429
Left engine shown. Right engine similar.
CS1_CS3_7321_018
MAX
Figure 9: Fuel Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-19
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
PROGNOSTICS AND HEALTH MANAGEMENT UNIT The prognostics and health management unit (PHMU) provides information about the health of the engine to help optimize maintenance operations. The PHMU is a single channel system with health monitoring functions using data from the EEC and engine sensors. The PHMU also performs the following functions: •
Engine vibration monitoring
•
Fan trim balance
•
Oil debris monitoring
•
Fan drive gear system (FDGS) auxiliary oil pressure sensor monitoring
The PHMU is powered from DC BUS 1. The PHMU communicates with the EEC through a CAN BUS. The engine health monitoring information is supplied to the HMU where is can be downloaded for analysis. The engine vibration monitoring helps locate sources of vibration and provides information for fan trim balance. The fan trim balance is carried out through the onboard maintenance system (OMS). The PHMU monitors the fan drive gear system (FDGS) to ensure that it receives a constant supply of oil under all flight conditions.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-20
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
EICAS DMC 1 and DMC 2
OMS HMU
DC BUS 1
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
ENGINE SENSORS
Engine Data
ELECTRONIC ENGINE CONTROL
P2/T2 Sensors Engine Vibration Oil Debris Monitoring Fan Drive Gear System Auxiliary Oil Pressure Sensor LEGEND Data Concentrator Unit Module Cabinet ARINC 429 CAN BUS
CS1_CS3_7320_004
DMC
Figure 10: Prognostics and Health Management Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-21
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
COMPONENT LOCATION
PERMANENT MAGNET ALTERNATOR GENERATOR
Components of the engine control system include:
The permanent magnet alternator generator (PMAG) is located on the main gearbox.
•
P2/T2 probe
•
Burner pressure (Pb) Sensor
•
Electronic engine control (EEC)
•
Prognostics and health management unit (PHMU)
•
Data storage unit (DSU)
•
Permanent magnet alternator generator (PMAG)
•
P2.5/T2.5 probe (Refer to Figure 12)
•
T3 probe (Refer to Figure 12)
P2/T2 PROBE The P2/T2 probe is mounted to the inlet cowl at the 11 o’clock position. BURNER PRESSURE SENSOR The Pb sensor is attached to the compressor intermediate case (CIC) fire seal at the 10 o’clock position. ELECTRONIC ENGINE CONTROL UNIT The EEC is located on the fan case at the 9 o’clock position. PROGNOSTICS AND HEALTH MANAGEMENT UNIT The PHMU is located beside the EEC on the engine fan case at the 9 o’clock position. DATA STORAGE UNIT The data storage unit (DSU) is located on the EEC.
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73-22
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
A
B BURNER PRESSURE (Pb) SENSOR
B C
A P2/T2 PROBE
Data Storage Unit (DSU)
Prognostic and Health Management Unit (PHMU)
D
C
PERMANENT MAGNET ALTERNATOR GENERATOR (PMAG)
CS1_CS3_7320_001
Electronic Engine Control (EEC)
D
Figure 11: Engine Control System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-23
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
P2.5/T2.5 PROBE The P2.5/T2.5 probe is located on the CIC at the 1 o’clock position. T3 PROBE The T3 probe is located on the diffuser case, forward of the fuel nozzles, at approximately the 1 o’clock position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-24
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
A B
B T3 PROBE
CS1_CS3_7320_003
A P2.5/T2.5 PROBE
Figure 12: P2.5/T2.5 Probe and T3 Probe Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-25
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
COMPONENT INFORMATION ELECTRONIC ENGINE CONTROL
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
The EEC electronic components are installed in a cast aluminum housing that is cooled by natural convection using fan compartment air.
The PHMU electronic components are installed in a cast aluminum housing that is cooled by natural convection using fan compartment air.
The EEC is installed on the fan case using four vibration isolators. The EEC has eight electrical connectors that allow it to interface with the aircraft systems and engine sensors, a data storage unit (DSU), and two pneumatic pressure ports.
The PHMU is installed on the fan case using four vibration isolators. The PHMU has two electrical connectors that allow it to interface with the EEC and engine sensors.
The EEC has an internal ambient pressure sensor that receives ambient pressure through a port on the lower right side of the EEC. The EEC also receives inlet pressure (P2) through the P2 port. DATA STORAGE UNIT The DSU is used for storage of engine configuration and identification data, engine historical and exceedance information, thrust ratings data, and fan trim balance data. If an EEC or prognostics and health management unit (PHMU) is replaced, the data stored in the DSU is supplied to the new unit. The DSU is mounted on the EEC A channel housing. The DSU is attached to the engine fan case by a lanyard and remains with the engine if the EEC is replaced.
CAUTION The DSU must remain with the engine when the EEC is removed. An incorrectly installed DSU can affect engine operation.
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73-26
CS130_73_L2_ControlSystem_ComponentInformation.fm
73 - Engine Fuel and Control 73-20 Control System
Electrical Connectors (8x)
Data Storage Unit Plug (DSU)
Vibration Isolators (4x)
P2 Sense Port
Electronic Engine Control (EEC)
Ambient Pressure Port
Electrical Connectors
Vibration Isolators (4x)
CS1_CS3_7321_006
Prognostics And Health Management Unit (PHMU)
Figure 13: Electronic Engine Control, Data Storage Unit, and Prognostics and Health Management Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-27
CS130_73_L2_ControlSystem_ComponentInformation.fm
73 - Engine Fuel and Control 73-20 Control System
PERMANENT MAGNET ALTERNATOR GENERATOR The permanent magnet alternator generator (PMAG) consists of a stator and a rotor. The stator and electrical windings are attached to the gearbox with captive bolts. The rotor remains attached to the gearbox shaft when the stator is removed. The PMAG stator has dual-channel EEC AC power windings, dual-channel N2 windings and a single generator winding for AC power to the fly-by-wire (FBW) controller.
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73-28
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73 - Engine Fuel and Control 73-20 Control System
Permanent Magnet Alternator Generator Stator
LEGEND Electronic Engine Control Channel A Power Electronic Engine Control Channel B Power Fly-By-Wire Power Speed (N2) Coils
CS1_CS3_7321_002
Permanent Magnet Alternator Generator Rotor
Figure 14: Permanent Magnet Alternator Generator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-29
CS130_73_L2_ControlSystem_ComponentInformation.fm
73 - Engine Fuel and Control 73-20 Control System
DETAILED COMPONENT INFORMATION ELECTRONIC ENGINE CONTROL POWER During normal operating conditions with the engine operating above 45% N2 speed, the EEC is powered by the permanent magnet alternator generator (PMAG). The PMAG provides a separate output for each EEC channel. When the engine is not running, both EEC channels are powered by the aircraft electrical system. Channel A of each EEC is powered by DC ESS BUS 3. Channel B of the left EEC is powered by DC ESS BUS 1. Channel B of the right EEC is powered by DC ESS BUS 2. When the engine is at or above minimum idle, a failure of one PMAG output results in the corresponding EEC channel switch its aircraft electrical power source. The other channel of the EEC continues operation using power from the PMAG.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-30
CS130_73_L3_ControlSystem_DetailedComponentInfo.fm
73 - Engine Fuel and Control 73-20 Control System
L PMAG
R PMAG
N2 is Greater Than 45% 34 VAC
N2 is Greater Than 45% 34 VAC
LEFT CBP
RIGHT CBP
CDC 1-5-1
CDC 2-5-1
L FADEC B DC ESS BUS 3
10
L FADEC A
DC ESS BUS 1
SSPC 10A
L FADEC B DC ESS BUS 3
10
R FADEC A
Logic: Always On
R EEC
Channel B
Channel A
Channel B
CS1_CS3_7321_014
LEGEND CBP CDC EEC PMAG
SSPC 10A
Logic: Always On
L EEC
Channel A
DC ESS BUS 2
Circuit Breaker Panel Control and Distribution Cabinet Electronic Engine Control Permanent Magnet Alternator Generator
Figure 15: Electronic Engine Control Power Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
ELECTRONIC ENGINE CONTROL FAULT DETECTION The EEC is designed to be fail-safe by detecting and accommodating component failures within the system. This is accomplished by the fault accommodation logic in the software. The main goal of the fault accommodation logic is to make the system operational upon the detection of all single electrical faults when the system is dispatched with two channels operational. The secondary goal is to maintain as much engine functionality as possible when multiple faults exist.
In the active-standby control scheme, the system gives active control to the healthiest channel. When the engine is shutdown, control of the engine switches to the other channel for the next engine start. This aids in the detection of potential hidden failures and provides equal output driver usage in the EEC. The channel switchover function is tested on each EEC power application.
Failures detectable by the control system are classified by their effect on the ability of the channel to control the engine. Many input failures have redundant backup sources and therefore do not require a channel switchover. The system is designed such that the minimum number of channel switchovers will occur due to faults. The fault accommodation is based on: •
If there are no detected faults in either channel, use average of the two channels
•
If an input fault or faults exist, use data from the other channel
•
If an output fault or faults exist, switch to other channel output
•
In the case of a dual failure, the EEC fails to a designated safe mode
The active-standby control scheme is used until the first failure is detected. At that point, control switches over to the active driver control scheme. The channel switchover logic has active driver and active-standby control schemes. The active driver control scheme allows either channel to control any of the output drivers independently, regardless of which channel is in control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
COMPONENT
ELECTRONIC ENGINE CONTROL
COMPONENT
Channel A
Channel A
Channel A
Channel A
Channel B
Channel B
Channel B
Channel B
NORMAL OPERATION
INPUT SWITCH
Both channels operating normally
Output from Channel A used
Channel A in control
Feedback from Channel B channels used
ELECTRONIC ENGINE CONTROL
COMPONENT
ELECTRONIC ENGINE CONTROL
COMPONENT
Channel A
Channel A
Channel A
Channel A
Channel B
Channel B
Channel B
Channel B
OUTPUT DRIVER SWITCH
CHANNEL SWITCH
Output from Channel B used
Channel A fault detected
Feedback from both channels used
Channel B in control
LEGEND Control Fault Feedback Standby
NOTE Channel A in control.
CS1_CS3_7321_027
ELECTRONIC ENGINE CONTROL
Figure 16: Electronic Engine Control Fault Detection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
PROGNOSTICS AND HEALTH MANAGEMENT UNIT The prognostics and health management unit (PHMU) performs engine health monitoring functions based on data provided by the EEC. The PHMU also performs the following functions: •
Engine vibration monitoring
•
Oil debris monitoring
•
Fan drive gear system (FDGS) oil monitoring
The PHMU provides engine health monitoring by comparing streaming engine data from engine sensors to stored models of engine gas path and subsystem parameters. The P2.5/T2.5 sensor provides P2.5 pressure to the PHMU. The other engine sensors provide information through the EEC. The information is processed by the PHMU and the appropriate output is sent to the EEC. The EEC then sends this information to the onboard maintenance system (OMS) and the health management unit (HMU). The PHMU receives signals from engine vibration sensors to assess the health of rotating and mechanical components. The vibration signals are compared against the engine speed signals supplied by the EEC to determine the location of the vibration. The processed vibration signals are sent to EICAS and also provide additional vibration information for fan balancing. The oil debris monitoring system checks for metal particles in the oil system and provides an early indication of engine wear. The PHMU supplies the information to the EEC as part of the data report supplied to the HMU. The PHMU also monitors the FDGS oil supply to ensure that a continuous supply of oil is being provided.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
CDC 1-10-8 L PHMU DC BUS 1
SSPC 5A
FWD VIBRATION SENSOR
P2.5/T2.5 PROBE
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
OIL DEBRIS MONITOR SENSOR
AFT VIBRATION SENSOR
AUXILIARY OIL PRESSURE SENSOR
Logic: Always On
EEC N1 Speed N2 Speed Fan Speed Engine Sensors
CDC EEC
Control and Distribution Cabinet Electronic Engine Control CAN BUS
NOTE Left engine shown. Right engine similar.
CS1_CS3_7321_015
LEGEND
Figure 17: Prognostics and Health Management Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
DATA STORAGE UNIT
Engine Thrust Rating
The digital storage unit (DSU) connects directly to EEC channel A. It communicates with both channels through a serial data bus. Data is stored in the DSU memory, EEC channel A, and EEC channel B.
The EEC uses the engine nameplate identification and aircraft type on the DSU to select which set of thrust ratings data to transfer to processor memory. During synchronization, the thrust ratings data corresponding to the synchronized engine nameplate identification and aircraft type are transferred to the EEC processor memory.
The DSU is capable of storing up to 8 MB of data. Data from the current flight overwrites older information when the memory section is full. The DSU stores the following information:
Engine Configuration Selector
•
Engine identification data
•
Trim balance data
•
Engine snapshot data
The engine configuration selector is used by EEC software to select control laws parameters compatible with the engine hardware configuration.
•
Engine historical data
DSU Failures
Engine configuration and identification data is programmed on the DSU during production acceptance testing and is not modifiable on the aircraft. The DSU remains with its specific engine for the life of the engine, unless the DSU fails.
If the DSU is not installed, and the aircraft is on the ground, the EEC does not allow the engine to start. If the DSU disconnects in flight, the EEC uses the programming information stored in the EEC memory to continue controlling the engine.
Engine identification data is read from the DSU during EEC initialization. Each channel performs an integrity check on the engine identification data in the DSU and in its own data flash. The DSU as the primary source, and uses its own data as a back-up source.
DSU critical faults such as nameplate and aircraft type incompatibility with the selectable engine ratings in one or both EEC channels result in an ENGINE FAULT advisory message and an 73 L(R) ENGINE FAULT FADEC FAULT 1 INFO message. The aircraft cannot be dispatched.
The engine serial number is used by the prognostics and health monitoring unit (PHMU) for engine condition monitoring.
A DSU noncritical fault such as a failure to write data to the DSU detected by one channel only results in an ENGINE FAULT advisory message and an 73 L(R) ENGINE FAULT - FADEC FAULT 2 INFO message.
N1 Modifier To ensure that all engines have uniform N1/thrust characteristics, an N1 class modifier is defined and its value is set in the DSU for each engine by tests during production. The EEC adds the N1 class modifier to the mechanical N1 speed to get the N1 feedback used in the N1 control loop and the N1 speed Indication displayed on EICAS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The DSU dispatch failures are only annunciated on the ground to prevent an ENGINE FAULT advisory message and associated INFO message from being displayed in flight should a DSU failure be detected after an in flight shutdown.
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73 - Engine Fuel and Control 73-20 Control System
DATA STORAGE UNIT
DATA STORAGE UNIT DATA Engine Identification Data
8 MB Flash
Trim Balance Data
DSU Hardware Part Number
Fan Weight Data
DSU Data Part Number
Low Pressure Turbine Weight Data
Engine Serial Number Trim Balance Solution Data Engine Nameplate Thrust (PW1525G, PW1524G, PW1521G, PW1521GA, or PW1519G) 8 MB Flash
8 MB Flash
Channel A Processor
Channel B Processor
Engine Snapshot Data
Engine Historical Data
Engine Faults (such as sensor or actuator failures)
Total Engine Running Time
Engine Events (such as exceedances and surges)
Start Counter
Engine Run Time in Flight
Shutdown Counter
Vibration History Data Aborted Start Counter Data for Trim Balance Calculations
Number of Flights
Aircraft Type (CS100, or CS300)
Preferred Channel in Control
N1 Modifier
Preferred Igniter
CRC32 Checksum
Next Overspeed Shutdown Test Set-up (channel A, channel B, or channel A and B)
ELECTRONIC ENGINE CONTROL
Overspeed Shutdown Test Failed Journal Oil Shuttle Valve Shutdown Test Failed
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
NVM Engine Position
CS1_CS3_7321_001
HPT Disk Temperature at Shutdown
LEGEND CAN BUS
Figure 18: Data Storage Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
CONTROLS AND INDICATIONS THROTTLE QUADRANT ASSEMBLY The throttle quadrant has two thrust levers to control the engines. Each thrust lever controls the forward and reverse thrust functions. The L(R) ENG switches provide engine run and shutdown commands. ENGINE PANEL The ENGINE panel CONT IGN PBA provides a manual start request signal to the EEC for engine starting. The START switch is spring loaded to the AUTO position for normal automatic starts. The switch has L(R) ENG CRANK positions for manual engine starts. ENGINE FIRE PANEL The ENGINE FIRE panel FIRE PBA provides a shutdown signal to the EEC when pressed. The FIRE PBA is used to shut down the engine should the L(R) ENG switch shutdown command fail.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
ENGINE CONT IGNITION
MAX
MAX
IDLE
IDLE
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
L ENG
R ENG
ON
ON
FF FF OFF
FF FF OFF
THROTTLE QUADRANT ASSEMBLY
L ENG
R ENG
F I RE FIRE E
F IRE IR E
BTL 1
BTL 2
BTL 1
BTL 2
AVAIL
AVAIL
AVAIL
AVAIL
L ENGINE FIRE PANEL
R ENGINE FIRE PANEL
CS1_CS3_7320_005
ENGINE PANEL
Figure 19: Fuel Control System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
INDICATIONS The fuel control system indications are displayed on the engine indication and crew alerting system (EICAS) page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
CLB 92.2
92.2
73.0
73.3
N1
VIB
718
722
EGT
86.5 VIB N2 86.6 1090 FF (KPH) 1110 124 OIL TEMP 124 116 OIL PRESS 116 FAN VIB 0.2 0.4 CS1_CS3_7320_008
EICAS
Figure 20: Fuel Control System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
DETAILED DESCRIPTION AIRCRAFT AND ELECTRONIC ENGINE CONTROL INTERFACE
Primary Flight Control Computers
The following aircraft systems interface with the EEC for engine control.
The primary flight control computers supply a steep approach signal used to modify the engine idle speed for a steep approach.
Flight Management System The flight management system (FMS) provides flight plan information used to calculate the engine thrust reference setting. The FMS also supplies the thrust mode selection automatically and is the interface for manual selection of thrust mode, automatic power reserve (APR) arming, N1 synchronization arming, and FLEX temperature selection.
Ice Detection The EEC uses the ice detector signal to control the cowl anti-ice valve (CAIV) operation. Engine Panel The manual or automatic engine start mode is selected from the ENGINE panel.
Integrated Air System Controllers The integrated air system controllers (IASCs) supplies the EEC with environmental control system (ECS) and wing anti-ice system configuration. The IASCs receive bleed status information from the EECs. Air Data Information The EEC uses aircraft air data as primary source of air data because of its accuracy and redundancy. The air data system supplies total air temperature, pressure, and Mach information. The air data system receives T2 information for use in the event of a TAT failure.
Autothrottle The EEC provides the N1 target command to the autothrottle system. The autothrottle controls the thrust levers based on the N1 target information. The autothrottle also provides engine trim signals for engine synchronization. ENGINE FIRE Panel The ENGINE FIRE panel supplies engine shutdown commands when an engine fire PBA is pressed.
Landing Gear and Steering Control Units
Throttle Quadrant Assembly
The EEC use the weight-on-wheels (WOW) signal for engine idle speed control, engine control, and thrust reverser control.
The throttle quadrant assembly (TQA) provides the EEC with thrust lever position information and engine ON/OFF commands. The EEC supplies a signal to energize the baulk solenoid when the thrust reverser translating sleeves are deploying or stowing.
Brake Data Concentrator Units The brake data concentrator units (BDCUs) supply wheel speed signals for control of the thrust reversers.
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73 - Engine Fuel and Control 73-20 Control System
FLIGHT MANAGEMENT SYSTEM Thrust Reference Flex Temperature APR Arming N1 SYNC Thrust Mode Selection
THROTTLE QUADRANT ASSEMBLY
Start Mode Thrust Lever Angle Engine ON/OFF
IASC 1 and IASC 2
Baulk Solenoid Command
ENGINE PANEL
ECS Configuration Wing Anti-ice Configuration Bleed Status
Autotrim N1 Target
AIR DATA SYSTEM
AUTOTHROTTLE
Total Air Temperature Pressure Mach DMC 1 and DMC 2
LGSCU 1 and LGSCU 2
ELECTRONIC ENGINE CONTROL
Weight-On-Wheels
BDCU 1 and BDCU 2 Ice Detected ICE DETECTOR PFCCs Steep Approach
Engine Shutdown Command LEGEND
ENGINE FIRE PANEL BDCU DMC IASC LGSCU PFCC
Brake Data Concentrator Unit Data Concentrator Unit Module Cabinet Integrated Air System Controller Landing Gear and Steering Control Unit Primary Flight Control Computers ARINC 429
CS1_CS3_7320_002
Wheel Speed
Figure 21: Aircraft and Electronic Engine Control Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
ENGINE SENSOR INTERFACE
Pamb Sensor
The engine sensors are used to provide fuel control and engine health monitoring.
The single-channel Pamb sensor measures the ambient pressure through a port on EEC. The Pamb sensor is connected to channel A of the EEC. Channel B receives Pamb data via the EEC CAN BUS.
P2/T2 Probe The P2/T2 probe consists of a single channel P2 probe that connects to channel B of the EEC and a dual channel T2 sensor. The P2 probe measures the engine inlet total pressure. P2 is used to validate the air data from the air data system, control scheduling functions, engine rating calculation and Mach number calculation. The P2 probe is heated. The left P2/T2 probe heater is powered by AC BUS 1 and the right P2/T2 is powered by AC BUS 2. The EEC commands the probe heat on when the engine is running above idle and the temperature is below 9°C (48°F). When the probe heat is on, the EEC software compensates for the effect of the probe heater. The EEC turns the P2/T2 heater off when the engine is shutdown or the temperature is above 9°C. If the EEC detects a heater failure, a 73 L ENGINE FAULT - P2/T2 HEATER INOP INFO message is displayed. The T2 sensor measures the engine total inlet temperature. The T2 sensor feeds both channels of the EEC. T2 is used to validate the total air temperature information from the air data system, control scheduling functions including engine rating calculation and Mach number calculation. The T2 sensor information is used in place of air data information when the aircraft is on the ground and below 60 kt. The P2T2 probe is used if air data from the air data system is not valid or not available. If the P2T2 probe also fails, the EEC uses default values: •
Inlet total temperature: synthesized total temperature based on standard day temperature, Mach and pressure
•
Total pressure: 14.7 psi
•
Pressure altitude: 17000 ft
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The engine sensed altitude is calculated based on Pamb. Pamb is used to validate the air data from the air data system, control scheduling functions, engine rating calculation and Mach number calculation. P2.5/T2.5 Sensor The P2.5/T2.5 sensor consists of a pressure probe and a temperature sensor. The P2.5 pressure probe measures the pressure at the low-pressure compressor (LPC) exit. The probe receives an excitation signal from the PHMU and provide an output used to monitor high-pressure compressor (HPC) and LPC health. The T2.5 temperature sensor measures total temperature at the LPC exit. The sensor receives excitation from channel A of the EEC and provides an output used to monitor high-pressure compressor (HPC) and LPC health. Channel B of the EEC receives T2.5 sensor data via the EEC CAN BUS. T3 Sensor The T3 thermocouple sensor provides a single output to channel B of the EEC. It measures the HPC exit temperature. T3 is used to assess engine performance and compressor health. Pb Sensor The Pb (burner pressure) sensor is a dual channel pressure sensor. The EEC provides excitation voltage to the sensor. Each channel has a low and a high range pressure transducer. Pb is used in fuel scheduling, stall detection, autostart logic, continuous ignition, and shaft shear detection.
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73 - Engine Fuel and Control 73-20 Control System
P2.5/T2.5 PROBE P2.5 T2.5
ELECTRONIC ENGINE CONTROL T3 SENSOR PROGNOSTICS AND HEALTH MANAGEMENT UNIT
Channel A
Channel B Pb SENSOR Channel A
PAMB
P2 Channel B
P2/T2 PROBE T2 CDC 1-15-7 L P2/T2 HTR AC SSPC 3A BUS 1
P2
LEGEND Pb
Air Pressure Burner Pressure CAN BUS
NOTE Left engine shown. Right engine similar.
CS1_CS3_7320_006
Logic: Single Input
Figure 22: Engine Sensor Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
ENGINE OPERATING LIMITS The engine operating limits ensure that the engine operates within the design parameters. The EEC uses feedback from engine sensors and inputs from aircraft systems to ensure that the design limits are not exceeded. A governing loop sets the primary engine thrust requirements using N1 as a reference. The governing loop is set based on aircraft system inputs for the thrust requirements. The N1 is limited to maximum speed of 10,600 rpm (100%) and a minimum of 1,616 rpm (15%). The minimum N1 limit is dependant on environmental conditions, minimum thrust level, ice accumulation and bleed air requirements. The governing loop is constrained by EEC limiting loops.
The fuel flow to burner pressure ratio limit loop makes sure that the fuel to air ratio inside the combustor is maintaining a margin against flameout. The burner pressure maximum limit loop protects the engine from over pressurization of the combustion chamber. The burner pressure minimum limit loop also ensures enough pressure for the environmental control system. The minimum burner pressure depends on the demands for engine air made by the following systems: •
Cowl anti-ice
•
Wing anti-ice
•
Environment control system
The fuel flow rate maximum and minimum limit loops are provided to prevent excessive fuel command changes during acceleration and deceleration. During engine starting, the EEC modifies the fuel schedule once the starter air valve (SAV) closes and the air turbine starter (ATS) disengages.
The following limiting loops are monitored by the EEC: •
N2
•
Fuel flow
•
Fuel flow/burner pressure ratio
•
Burner pressure
Thrust Limitation at Low Airspeed
•
Fuel flow rate
A thrust limitation at low air speed (TLLS) function is available in TO mode only. TLLS maintains a minimum nose landing gear loading during the takeoff phase.
The N2 is limited to maximum speed of 24,470 rpm (100%) and a minimum of 13,087 rpm (53%). The minimum N2 limit ensures that the engine remains above generator trip speed. If N1 or N2 increase beyond 100%, but do not reach the overspeed shutdown condition of 105%, a L(R) ENG EXCEEDANCE caution message is displayed and the engine thrust must be reduced manually using the thrust levers. The fuel flow maximum limit protects the IFPC. There is a fuel flow minimum that assures a minimum amount of fuel is delivered to the combustor during starts. While in flight, the minimum fuel flow provides flameout protection at high altitude and during high G maneuvers. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The TLLS function is enabled when the aircraft is on the ground and the airspeed is less than 80 kt. As the airspeed increases, the TLLS input is reduced and the required thrust commanded by the EEC is provided by 80 kt. The TLLS function is inhibited when the aircraft is on the ground and the parking brake is set. This ensures full power can be developed during a high-power ground run.
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73 - Engine Fuel and Control 73-20 Control System
ENVIRONMENT CONTROL SYSTEM
DMC 1 and DMC 2
AIRCRAFT SYSTEMS
WING ANTI-ICE SYSTEM
EEC N1 SPEED SENSOR
PMAG (N2)
Governing Loop • N1 Limiting Loops • N2 • Fuel Flow • Fuel Flow/Burner Pressure Ratio • Burner Pressure • Fuel Flow Rate
BURNER PRESSURE SENSOR
COWL ANTI-ICE
STARTER AIR VALVE
AIR TURBINE STARTER
IFPC DMC EEC IFPC PMAG
Data Concentrator Unit Module Cabinet Electronic Engine Control Integrated Fuel Pump and Control Permanent Magnet Alternator Generator ARINC 429
Fuel Metering Valve NOTE LH Engine shown, RH similar.
CS1_CS3_7321_033
LEGEND
Figure 23: Electronic Engine Control Engine Operating Limits Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
OVERSPEED PROTECTION, SHAFT SHEAR PROTECTION AND THRUST CONTROL MALFUNCTION Overspeed protection
Thrust Control Malfunction
Overspeed protection is provided for both high (N2) and low (N1) pressure rotors. This is done via dedicated protection computers within the EEC independent of the channel A and B control computers.
A thrust control malfunction is an uncontrollable high thrust event. A thrust control malfunction is characterized by:
The protection computers control the overspeed shutdown solenoid (OSS). When either N1 or N2 engine speeds exceed the overspeed limit, at any time the engine is running, the engine immediately shuts down.
•
Uncommanded engine thrust increase
•
Engine thrust remaining at a high thrust setting when a lower thrust setting is commanded
The protection computers share the N1 and N2 signals provided to the control computers within the EEC and have a separate crosstalk CAN BUS.
In either case, the thrust can not be controlled by the thrust lever. When the EEC detects a thrust control malfunction, the OSS is energized to shut down the engine
The EEC limits fuel flow when redline limits are reached for N1 or N2. When either of the protection computer detects an overspeed limit of 105% N1 or 105% N2, the OSS is energized.
Thrust control malfunction detection is inhibited during:
Once an overspeed limit exceedance is detected, the system is latched and must be reset. The system can be reset by cycling the power supply to the EEC, or by selecting the L(R) ENG switch to ON to initiate the engine start sequence. To ensure that there are no undetected faults, the OSS is tested at shutdown.
•
Thrust reverse operation
•
Approach
•
Steep approach
Shaft Shear Protection The protection computers also provide protection against low and high pressure rotor shaft shear. If the protection computers determine that either shaft shear conditions have occurred, the OSS is energized to shutdown the engine.
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73 - Engine Fuel and Control 73-20 Control System
L ENG
R ENG
ON
ON
OFF
OFF
THROTTLE QUADRANT ASSEMBLY
INTEGRATED FUEL PUMP AND CONTROL
ELECTRONIC ENGINE CONTROL Channel A
N1 SPEED SENSOR
Control Computer
PERMANENT MAGNET ALTERNATOR GENERATOR
Overspeed Shutdown Solenoid
Channel B
Protection Computer
CS1_CS3_7321_034
Control Computer
N2
Protection Computer
LEGEND CAN BUS
Figure 24: Overspeed, Shaft Shear Protection, and Thrust Control Malfunction Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL OPERATION - START The IFPC is controlled by the EEC to provide metered fuel for combustion based on thrust lever position and ambient operating conditions. The IFPC also provides servo fuel for the fuel actuated system on the engine. The IFPC consists of both hydromechanical and electromechanical valves to control the fuel flow. The fuel metering valve (FMV) is a dual channel component controlled by the EEC. A dual channel torque motor allows either channel of the EEC to control the FMV. FMV feedback is provided to the EEC using a dual coil LVDT. The pressure-regulating valve (PRV) maintains a constant pressure drop across the FMV by returning excess pump flow to the inlet of the high-pressure pump. The excess flow is dependent on the servo fuel demand and metered fuel requirements. The fuel flow meter provides the EEC with an indication of the amount of fuel being used for combustion. When the engine is commanded to shutdown, the windmill bypass valve (WBV) opens to connect fuel from upstream of the minimum pressureshutoff valve (MPSOV) with the fuel filter inlet. The minimum pressure-shutoff valve (MPSOV) is a spring-loaded valve that prevents fuel from entering the primary and secondary fuel manifolds until it is sufficiently pressurized. The overspeed shutdown solenoid (OSS) is a dual-coil solenoid valve that operates the WBV and MPSOV to shutoff metered flow to the engine. The flow divider valve (FDV) schedules the fuel flow to primary and secondary manifolds depending on engine start requirements. The FDV is normally spring-loaded to the closed position, blocking fuel to the secondary fuel manifold. Once the fuel pressure between the primary and secondary fuel manifolds exceeds 100 psi, the FDV opens allowing fuel into the secondary manifold.
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73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL From Servo Pressure Regulator MAX
MAX
IDLE
IDLE
To Fuel Filter
OSS
Pressure-Regulating Valve
SOL
WBV L ENG
R ENG
ON
ON
OFF
OFF
FMV
TQA
Channel A
From Main Pump
EEC
Channel B Channel A
T/M SOL
EBV SOLENOID Channel B MPSOV
SOL T/M
EBV
FUEL/OIL MANIFOLD
To Primary Fuel Manifold To Secondary Fuel Manifold FUEL FLOW METER CS1_CS3_7321_039
EBV EEC FDV FMV MPSOV OSS TQA WBV
Metered Fuel Pressurized Fuel Return Fuel Servo Fuel Equalization Bypass Valve Electronic Engine Control Flow Divider Valve Fuel Metering Valve Minimum Pressure-Shutoff Valve Overspeed Shutdown Solenoid Throttle Quadrant Assembly Windmill Bypass Valve CAN BUS Solenoid Torque Motor
FDV
Figure 25: Integrated Fuel Pump and Control Operation - Engine Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL OPERATION - RUN The EEC energizes the EBV solenoid when the engine start is completed. The equalization bypass valve solenoid controls the equalization bypass valve (EBV). The EBV energizes to equalize the fuel flow in the primary and secondary fuel manifolds. The FDV closes and the secondary fuel manifold is supplied through the EBV.
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73-52
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL From Servo Pressure Regulator MAX
MAX
IDLE
IDLE
To Fuel Filter
OSS
Pressure-Regulating Valve
SOL
WBV L ENG
R ENG
ON
ON
OFF
OFF
FMV
TQA
Channel A
From Main Pump
EEC
Channel B Channel A
T/M SOL
EBV SOLENOID Channel B MPSOV
LEGEND
SOL T/M
EBV
FUEL/OIL MANIFOLD
To Primary Fuel Manifold To Secondary Fuel Manifold FUEL FLOW METER CS1_CS3_7321_008
EBV EEC FDV FMV MPSOV OSS TQA WBV
Metered Fuel Pressurized Fuel Return Fuel Servo Fuel Equalization Bypass Valve Electronic Engine Control Flow Divider Valve Fuel Metering Valve Minimum Pressure-Shutoff Valve Overspeed Shutdown Solenoid Throttle Quadrant Assembly Windmill Bypass Valve CAN BUS Solenoid Torque Motor
FDV
Figure 26: Integrated Fuel Pump and Control Operation - Engine Run Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-53
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
NORMAL ENGINE SHUTDOWN On Ground
In Flight
The EEC commands shutdown of the engine using the overspeed shutdown solenoid (OSS) whenever one of the following conditions is present:
In flight the EEC commands the engine to shutdown by driving the FMV to the minimum flow position first to allow time for the high-pressure compressor (HPC) stator vane actuator to track to the optimal start position.
•
L(R) ENG SWITCH is moved to the OFF position on the throttle quadrant assembly
•
An excessive engine overspeed is detected by the EEC
•
A shaft shear is detected by the EEC
When the HPC stator vane actuator is in position, the fuel metering valve is commanded closed and the OSS is energized.
The EEC uses the overspeed protection system to shut down the engine. The EEC channel in control commands the overspeed shutoff solenoid to energize. On every third shutdown both EEC channels command the overspeed shutoff solenoid to fully test the system. When the EEC energizes the OSS, servo fuel pressure opens the WBV to allow fuel to flow back to the pressure-regulating valve. The pressureregulating valve opens so that the fuel flows back to the fuel filter and is recirculated. At the same time, the MPSOV moves to the closed position and stops fuel from flowing to the engine fuel nozzles. The EEC de-energizes the equalization bypass valve solenoid to the closed position. Once the overspeed protection system has shut down the engine, or if the overspeed protection system fails to shut down the engine, the EEC commands the fuel metering valve closed. When the fuel metering valve is confirmed closed, the overspeed shutdown solenoid is de-energized and all EEC processors are reset.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-54
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL From Servo Pressure Regulator MAX
MAX
IDLE
IDLE
To Fuel Filter
OSS
Pressure-Regulating Valve
SOL
WBV L ENG
R ENG
ON
ON
FF FF OFF
FF FF OFF
FMV
TQA
Channel A
From Main Pump
EEC
Channel B Channel A
T/M SOL
EBV SOLENOID Channel B MPSOV
SOL T/M
FUEL/OIL MANIFOLD
EBV
To Primary Fuel Manifold To Secondary Fuel Manifold FUEL FLOW METER CS1_CS3_7321_038
EBV EEC FDV FMV MPSOV OSS TQA WBV
Metered Fuel Pressurized Fuel Return Fuel Servo Fuel Equalization Bypass Valve Electronic Engine Control Flow Divider Valve Fuel Metering Valve Minimum Pressure-Shutoff Valve Overspeed Shutdown Solenoid Throttle Quadrant Assembly Windmill Bypass Valve CAN BUS Solenoid Torque Motor
FDV
Figure 27: Normal Engine Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-55
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
EMERGENCY SHUTDOWN An emergency shutdown can be carried out using the L ENG or R ENG FIRE PBA. Pressing the fire PBA initiates the following: •
Engine feed shutoff valve closes
•
Pressure-regulating shutoff valve (PRSOV) closes
•
Fan air valve closes
•
Hydraulic shutoff valve closes
•
Generator trips offline
•
EEC commands engine shutdown
•
Both engine fire extinguisher bottles are armed
The integrated air system controller (IASC) and electronic engine control (EEC) receive digital inputs from the FIRE PBA via an ARINC 429 BUS. The fan air valve and pressure-regulating shutoff valves are closed by the IASC. The EEC provides a shutdown command to the engine. The engine feed shutoff valve is closed through switch contacts in the FIRE PBA.
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73-56
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
INTEGRATED COCKPIT CONTROL PANEL Left Engine Fire Panel BTL 1
L ENG
AVAIL FIRE
EPC 1 BATT DIR BUS 1
5
Firex A
To Fire Extinguisher 1
EPC 2 BATT DIR BUS 2
5
BTL 2
Firex B
K1
AVAIL
To Fire Extinguisher 2 From L GEN PBA
K2
L Generator Offline
NOTE
PIM
Left engine shown. Right engine similar. LEGEND Data Concentrator Unit Module Cabinet Electronic Engine Control Electrical Power Center Integrated Air System Controller Panel Interface Module ARINC 429
L IASC
L EEC
LEFT ENGINE FEED SHUTOFF Closed VALVE Open
HYD SOV Closed
CS1_CS3_7321_031
DMC EEC EPC IASC PIM
DMC 1 and DMC 2
Figure 28: Emergency Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-57
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
ENGINE POWER MANAGEMENT Engine Bleed Management Each EEC receives information from the integrated air system controllers (IASCs). IASC 1 provides information to the left EEC. IASC 2 provides information to the right EEC. •
High-pressure valve (HPV) open or closed status
•
Environmental control system (ECS) configuration
•
Anti-ice configuration: -
Wing anti-ice status
-
Nacelle anti-ice status
Idle Speed The EEC selects engine idle speed based on engine requirements and bleed demand. The minimum idle speed is based on the following priorities: •
Minimum fuel flow for combustor stability
•
Minimum pressure to provide the required aircraft bleed flow, and maintaining the minimum pressure to not exceed a 20% core flow bleed limit
•
Minimum N1 to achieve acceleration time requirements
•
Minimum N2 speed to maintain the PMAG power generation
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
IASC INPUTS
High-Pressure Bleed Port • Closed • Open
EEC IDLE SETTING DECISION TREE
1. Minimum FUEL flow for combustor stability.
2. Minimum pressure to provide aircraft bleed flow (not to exceed 20% of the core flow).
Select Highest Value
Set Idle
ECS Bleeds • No Flow • Low-Flow • High-Flow 3. N1 Idle for transient acceleration times.
Wing Anti-Ice • No Flow • Low-Flow • High-Flow
LEGEND Environmental Control System Electronic Engine Control Integrated Air System Controller
CS1_CS3_7321_029
ECS EEC IASC
4. N2 Idle to maintain minimum generator speed.
Figure 29: Engine Bleed Management and Idle Speed Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-59
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
ENGINE POWER SETTINGS AND MODES Using flight plan performance data from the flight management system (FMS), environmental data, and engine sensors, the EEC calculates the N1 target for maximum takeoff and go-around, maximum continuous thrust, derated takeoff, reduced takeoff (FLEX), and climb thrust. Maximum Thrust The EEC sets the maximum thrust available at the MAX thrust position to maximum takeoff thrust (MTO) when the aircraft is operating in the takeoff or go around envelope. Outside the takeoff envelope or go around envelope, the EEC schedules maximum continuous thrust (MCT). The MCT rating supports one engine inoperative (OEI) conditions and is designed to be lower or equal to MTO thrust and higher than or equal to climb thrust (CLB). Mode Selection The thrust lever is adjusted automatically to the selected thrust mode rating when the autothrottle is engaged. Thrust mode changes are managed automatically based on the FMS active flight plan. The thrust mode can be controlled manually on the FMS PERF page DEP tab using the THRUST button. A window pops up to enable manual selection. Derated Takeoff Derated takeoffs TO-1, TO-2, or TO-3 are used to reduce fuel consumption, noise, and to extend engine life and reliability. The derated TO thrust selection is computed using the mode selection on the FMS DEP page. The takeoff derate is limited to a minimum of 17,000 pounds of thrust.
the assumed temperature input is higher than the actual temperature, the FLEX reduction in thrust is applied for that takeoff. FLEX is available from any selected takeoff thrust mode. Climb Thrust The FMS calculates the optimal climb (CLB) thrust and N1 reference based on the takeoff thrust mode (derated, FLEX), and the outside air temperature (OAT). Derated climb CLB-1 or CLB-2 allow the thrust settings for the climb envelope of the flight to be reduced by the FMS. There are two levels of climb derate. Each subsequent level reduces the thrust during the climb. Go-Around Go-around thrust is set as a thrust reference when the aircraft is on approach with the slats and flaps extended and the landing gear down. Go-around thrust is equivalent to MTO. Go-around thrust is initiated when the aircraft is in flight and the takeoff/go-around (TOGA) switches on the thrust levers are pressed. Idle Speed The EEC determines the idle setting based on engine stability requirements, bleed demand, and phase of flight. The minimum idle speed setting minimizes the aircraft landing distance and brake wear during taxi. Flight idle is used in flight to minimize fuel consumption during descent. When the landing gear is selected down or the flaps are extended, the engine idle speed increases to the approach idle setting to allow a faster engine response during a go-around. During a steep approach, a lower than normal approach idle setting is used. The steep approach idle is adjusted for bleed air demands and goaround thrust requirements.
FLEX Thrust FLEX thrust is computed using the assumed temperature method, based on the value entered in the FMS. On the ground, when the EEC confirms Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-60
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
Departure Tab
ACT
FMS1
POS
CLB
DEP Flex
DBASEE DBAS
FPLN
CR CRZ RZ
DES
ORIGIN
RWY
SID
LFBO
RW06L
CHATY2 ------
TO THRUST
FLEX
OAT
TO-1
+15
°C
+10
PERF
ROUTE ROUT UTE ARRR
RWY WIND °C
TO-3
Performance Page Selection
TRANS
FLEX 44°C
CLB THRUST
080°// 14
73.3
73.3
73.3
73.3
CLB-1 N1
APR
ARMED Fixed Takeoff Derate Selection
V1
EICAS
Fixed Climb Derate Selection
V2
SET VSPEEDS DS
123 125 130
APR Selection
COMPLETE F3
3
150
BLEEDTHRUST BLEE SOURCE
APU AP
F2 160
F1 170
180
WING ANTI-ICE
MODE TRANS
CLB THR CLB-1 CLB-2
Symbol
GD 210
Manually entered thrust reference value.
OFF OPT ALT MAX ALT STAB TRIM FL330 FL410 ND 6.2
PITCH SPD THRUST OEI FMS PERF VFLC FMS CLB TO
80 %N1 VFLC RED FLAPS CLEAN %N1
Description FMS managed thrust reference value.
COWL ANTI-ICE
OFF
CRZ ALT AUTO TOW TOW CG FL250 5500 LB AUTO 55000 33.5 %MAC TO-1 + APR PR 100 %N1 DEP PROFILE MAN HT/ALT 100 %N1 GA / 1618 STANDARD MCT ACCEL 1500 94 ST %N1 Thrust Button
F0
CNCL C NC L NCL
FMS CLB CLB
ENGINE SYNC
%N1
DONE D DO O NE
THRUST...
Thrust mode. TO-3 Assumed temperature for FLEX XXºC. FLEX 44°C 73.3 73.3 Digital thrust reference.
TO TO-1 TO-2 TO-3 CLB CLB-1 CLB-2 MCT GA
or or or or or or or or or
TO TO-1 TO-2 TO-3 CLB CLB-1 Magenta: FMS managed CLB-2 Cyan: manual selection MCT GA
FLEX 44°C 73.3 73.3 FLEX 44°C 73.3 73.3
FMS PERFORMANCE PAGE
CS1_CS3_7321_036
SLAT/FLAP Auto Manual Selection
VR
Figure 30: Engine Power Settings and Modes Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-61
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
AUTOMATIC POWER RESERVE
Scenario 3
The automatic power reserve (APR) automatically provides a higher thrust setting to the operating engine should one of the engines fail during a derated takeoff. APR provides improved single-engine acceleration and climb gradient performance.
The engines are derated to TO-3 and FLEX is selected. When the EEC detects one engine has lost thrust, the derate changes to TO-2 and FLEX is canceled. The operative engine is commanded to supply TO-2 thrust.
APR is only available when a derated takeoff thrust reference is selected. APR is not available for any other thrust mode. Should one engine fail, the EEC automatically increases the thrust on the operative engine by canceling the engine fixed derate by one level and removing any FLEX entry. When the thrust increase is commanded, the thrust levers angle (TLA) does not change.
Additional thrust is available by advancing the thrust levers. The EEC recalculates the TLA versus N1 command relationship to ensure the full range of thrust is available from the engine.
If TO-1 derate is selected, the engine thrust increases to MTO. If TO-2 or TO-3 are selected, the thrust is commanded to a level below MTO in order to maintain aircraft controllability. MTO is available by advancing the thrust levers to MAX. When armed, APR activates when both thrust levers are advanced beyond 23° and one engine experiences a loss of power greater than 15% N1. The APR can be manually disarmed from the FMS performance page DEP tab. When APR is manually disarmed selected, an APR DISARM status message is displayed. Scenario 1 The engines are derated to TO-1 prior to takeoff. This cancels one fixed derate level and maximum thrust is commanded on the operative engine. Scenario 2 The engines are derated to TO-1 and FLEX is selected. When the EEC detects one engine has lost thrust, the derate and FLEX are canceled. Maximum thrust is commanded on the operative engine.
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TECHNICAL TRAINING MANUAL
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
LEGEND APR MTO TLA TO
Thrust
Automatic Power Reserve Maximum Takeoff Thrust Lever Angle Takeoff
B2
SCENARIOS
A2
23,300 lbf
Scenario 1
A1 to A2
The derate is canceled and thrust is increased to MTO.
A1
21,000 lbf
Scenario 2
B1 to B2
Both the derate and the FLEX are canceled, thrust is increased to MTO.
C1 to C2
The derate is reduced to TO-2 and FLEX are canceled. Engine thrust automatically increases to TO-2. Additional thrust is available by advancing the thrust levers.
TO-1
18,900 lbf
C2
Scenario 3
17,000 lbf APR Activated
B1 C1
TO-2
TO-1 + FLEX
TO-3 + FLEX
88.5
APR 88.5
20.0
88.5
N1 EICAS PAGE
TO THRUST
0° TO-3
TO-2
TO-1
TO (47.5°)
TO-1
+15
°C
APR
ARMED
THRUST LEVER ANGLE VERSUS APR RELATIONSHIP
FMS PERFORMANCE PAGE
CS1_CS3_7321_030
Thrust Lever Angle
FLEX
Figure 31: Automatic Power Reserve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
N1 SYNCHRONIZATION The N1 synchronization (N1 SYNC) function is used to synchronize the engine speeds to reduce cabin noise. The N1 SYNC function is selectable through the aircraft flight management system (FMS) and is armed by default. The N1 SYNC function maintains the difference between the sensed N1 values of both engines to less than 0.2% during steady state conditions. N1 SYNC is enabled by the EEC, provided specific engine operating conditions are satisfied. The left engine is the master engine for the N1 SYNC function. The N1 command for the slave engine is biased to match that of the master engine. With the N1 SYNC mode set in the FMS, the EEC activates N1 SYNC if the following conditions are met: •
Both engines are operating in a steady-state
•
The aircraft is not in takeoff, or go-around
•
The difference between the two engines N1 commands is less than 1.5% N1
•
Throttle lever angle (TLA) is greater than 6°
•
No opposite engine low-thrust condition or automatic power reserve selection is detected
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73 - Engine Fuel and Control 73-20 Control System
THRUST MODE TRANS CRZ ALT FL250
AUTO AUTO TO-1 + APR MAN GA
MCT CLB CLB-1 CLB-2
OPT ALT FL330
MAX ALT FL410
100 %N1 100 %N1 94 %N1
OEI FMS PERF
CCNCL NCL NCL
80 %N1 %N1
ENGINE SYNC
%N1
DONE D DO ONE ONE FMS PERFORMANCE PAGE
DATA CONCENTRATOR UNIT MODULE CABINET
N1 SYNC Command
Electronic Engine Control
92.2
92.2
73.0
73.2
N1 SYNC EICAS
Integrated Fuel Pump And Control LEFT ENGINE
Integrated Fuel Pump And Control
LEGEND ARINC 429
RIGHT ENGINE
CS1_CS3_7321_028
Electronic Engine Control
CLB FLEX 44°C
Figure 32: N1 Synchronization Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
FAULT REPORTING AND TESTING INTERFACE
-
No dispatch
Each EEC channel has a non-volatile memory (NVM) that can store snapshots and transient records of parameters. Each EEC channel records its own faults or events that are detected. Old data is overwritten by new data when the memory is full.
-
Short time dispatch
-
Long time dispatch
The PHMU transmits data to the health management unit (HMU) via the EEC. The EEC sends the data via an ARINC 429 bus for distribution to the health management unit (HMU) and the onboard maintenance system (OMS). The EEC also provides information to generate CAS and INFO messages on EICAS and the INFO synoptic pages. The CAS and INFO messages are used to determine the capability to dispatch the aircraft. The health management unit is capable of recording the following: •
Fault data such as failures detected within FADEC system components or its interfaces with the aircraft such as sensor and actuator failures
•
Event data consisting of abnormal operation such as surges and exhaust gas temperature exceedances
•
Health trending data consisting of parameters collected on every flight to track engine performance over time. The data is used for trend monitoring and can provide early detection of an engine failure. The data can also be used to generate reports such as takeoff reports and stable cruise reports
The OMS records faults annunciated by the EEC. The OMS also provides the capability of testing EEC functions and displaying real-time data that support line maintenance operations. (see ATA 45). Dispatch The EEC monitors all faults and classifies them according to the operability of the engine. Faults are classified as follows: •
•
Master minimum equipment list (MMEL) faults
Time limited dispatch faults include: •
Failures directly used for thrust control
•
Malfunctions resulting in an auto-idle or auto-shutdown condition
•
Malfunctions resulting in a thrust lever moved to idle or an engine shutdown in response to a CAS message
No dispatch faults are major failures that prevent operation or control of the engine. A no dispatch fault has an L(R) ENGINE FAULT advisory message and a 73 L (R) ENGINE FAULT - FADEC FAULT 1 INFO message. Short time dispatch faults usually relate to a single channel failure that can be compensated for by alternate means on the engine. Short time dispatch faults are limited to 125 flight hours. A short time fault has an L(R) ENGINE FAULT advisory message and a 73 L (R) ENGINE FAULT - FADEC FAULT 2 INFO message. Long time dispatch faults usually relate to a failure that does not affect the operability of the engine such as a sensor dedicated to engine performance monitoring. Long time dispatch faults are limited to 425 flight hours. A long time fault has an L(R) ENGINE FAULT advisory message and a 73 L (R) ENGINE FAULT - FADEC FAULT 3 INFO message. MMEL faults are normally related to secondary systems or indication system not essential for engine operation. MMEL faults are generally associated with caution messages.
Time limited dispatch
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73 - Engine Fuel and Control 73-20 Control System
Maintenance Main Menu
DMC 1 and DMC 2
View Post Flight Summary View Flight Deck Effects ( 28 Active) View Fault Messages ( 43 Active) View Service Messages ( 13 Active) View System Exceedances ( 0
Stored)
View System Trends
HEALTH MANAGEMENT UNIT
View Aircraft Life Cycle Data View System Parameters Perform LRU/System Operations
ELECTRONIC ENGINE CONTROL
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
R ENG OPER DEGRADED R ENGINE FAULT XXXXXXXXXXXXXXXXXXX XXXXXXXXXXXXXXXXXXX
View System Configuration Data Write Maintenance Reports
Utility Functions
OMS
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
73
R ENGINE FAULT - FADEC FAULT 2
30
L ENG ANTI ICE - LOW PRESS SW INOP
LEGEND DMC OMS
Data Concentrator Unit Module Cabinet Onboard Maintenance System ARINC 429 CAN BUS
PREVIOUSLY ACKNOWLEDGED 27
FLIGHT CONTROL FAULT - AILERON FORCE MON INOP
SYNOPTIC PAGE – INFO
CS1_CS3_7320_007
EICAS
Figure 33: Electronic Engine Control Fault Reporting and Testing Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine controls system.
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CS130_73_L2_ControlSystem_MonitoringTests.fm
73 - Engine Fuel and Control 73-20 Control System
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L (R) ENG EXCEEDANCE Left (R) N1 or left (R) N2 or left ITT or left oil temperature above threshold. L (R) ENG OPER DEGRADED
Uncertainty on T2/TAT (degraded or default value), or HPC stator vane actuator not tracking as it should.
L (R) ENG FAIL
L 9R) engine Sub-idle.
ENG SETTINGMISMATCH
Thrust management data received or fed back by FADEC not matching the APR setting, N1 target, etc.
Table 2: ADVISORY Messages MESSAGE L (R) ENGINE FAULT
LOGIC Loss of redundant or non-critical function for the left engine.
Table 3: STATUS Messages MESSAGE APR DISARMED
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LOGIC APR is disarmed in the FMS and a derated takeoff is being used.
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
Page Intentionally Left Blank
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73 - Engine Fuel and Control 73-20 Control System
Table 4: INFO Messages Table 4: INFO Messages MESSAGE
MESSAGE LOGIC
73 L (R) ENGINE FAULT - EEC A CTRL CPU INOP
EEC channel A control processor has failed.
73 L (R) ENGINE FAULT - EEC B CTRL CPU INOP
EEC channel B control processor has failed.
73 L (R) ENGINE FAULT - PT/T2 HEATER INOP
FADEC logic has determined that the P2/T2 heater has a logical failure and is indicated as failed.
73 L (R) ENGINE FAULT - FADEC FAULT 1
Collector bit for all non-dispatchable FADEC faults.
73 L (R) ENGINE FAULT - FADEC FAULT 2
Collector bit for all FADEC short term TLD faults.
73 L (R) ENGINE FAULT - FADEC FAULT 3
Collector bit for all FADEC long term TLD faults.
77 L (R) ENGINE FAULT - PHMU INOP
FADEC logic has determined that the PHMU is failed.
73 L (R) ENGINE FAULT - EEC 28VDC REDUND LOSS
Set when only one EEC channel of the left (right) engine has detected loss of 28 VDC input voltage.
73 L (R) ENGINE FAULT - HEALTH MON DEGRADED
Set by the left (right) engine when the P2.5 or T2.5 sensor is failed.
73 L (R) ENGINE FAULT - T3 SNSR INOP
Set by the left (right) engine when the T3 sensor is failed. This INFO message will be masked if the EEC Channel B control processor is failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
73 L (R) ENGINE FAULT - FUEL EQUAL INOP
TECHNICAL TRAINING MANUAL
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LOGIC Set by the left (right) engine in any of the following conditions: Both EEC channels are unable to command fuel equalization solenoid. Any channel control processor failed and remaining channel unable to command fuel equalization solenoid.
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73 - Engine Fuel and Control 73-20 Control System
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM
System Interactive Test
The aircraft onboard maintenance system (OMS) uses the electronic engine control (EEC) to perform functional checks of specific systems. Once the new EEC is installed, the EEC SYSTEM TEST must be carried out. This test is performed through the OMS using one channel at a time. Each system test should be checked through both channels.
The system interactive test procedure is used to check for satisfactory operation of the EEC system before an engine start.
The following EEC tests can be performed through the aircraft OMS:
This procedure is used to verify the presence of existing faults, and to confirm that corrective action has eliminated the fault. The EEC system interactive test electrically verifies the EEC, and prognostics and health management unit (PHMU) input and output circuits.
•
Actuator Test
•
EEC System Test
•
HPD Test
•
Harness Test
•
Ignition Test
•
P2 Probe Heat Test
Harness Interactive Test
•
PCE Door Test
•
Thrust Reverser Cycling Test
The harness interactive test checks for harness and connector failures. During the test suspected harnesses must be moved to help reproduce the fault messages reported by the EEC.
NOTE The system interactive test does not test igniters, thrust reverser, or hydraulic pumps. Separate interactive tests are provided for these systems.
This procedure is used to perform harness checks in order to isolate intermittent EEC monitored faults. Probe Heat Test The P2/T2 probe heat test verifies the operation of the probe heater circuit. The EEC turns the probe heat on, and after 60 seconds verifies a rise in T2 temperature. The EEC then turns the probe heat off.
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73-72
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-20 Control System
MAINT MENU LRU/System Operations View
All ATAs
Select ATA
73 - ENGINE FUEL & CONTROL
Select LRU/System Left Engine EEC (1) channel A
DATA
RIGGING
TEST
NVM
Actuator Test EEC System Test HPD Test Harness Test Ignition Test P2 Probe Heater Test PCE Door Test
CS1_CS3_7330_003
Thrust Reverser Cycling Test
Figure 34: OMS EEC System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
73-73
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-20 Control System
HMU Power Assurance Test The power assurance test can be performed using the ATA 46 HMU CH A or CH B selection on the OMS LRU/System Operations page. Using ambient pressure, temperature, N1 errors, and nacelle anti-ice (NAI) status, the HMU software calculates and corrects the N2, EGT, and fuel flow parameter range. During the engine run, the HMU compares the actual engine parameters against the corrected parameter ranges. If all of the actual parameters are all within the corrected parameter ranges, a TEST PASSED message is displayed. A TEST FAIL message is shown if any parameter is out of range.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-74
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-20 Control System
MAINT MENU
MAINT MENU EMU-Engine 1 Power Assurance Test
LRU/System Operations View
All ATAs
Select ATA
46 - INFORMATION SYSTEMS 1/6 *** WARNING *** BOTH ENGINES WILL OPERATE DURING THE TEST. ENSURE TO FOLLOW THE ENGINE RUN-UP SAFETY PRECAUTIONS
Select LRU/System HMU CH A (A604)
DATA
TEST
NVM
*NOTE* TEST DURATION ~ 20 MINUTES
Engine 1 Power Assurance Test HMU - Cellular Connectivity Test HMU - IMS Interface Test
PRECONDITIONS
HMU - NVM Erase Function
1- REFER TO AMP XX-XX-XX-XX-XXXX 2- MAKE SURE THAT WIND SPEED < 35 KNOTS 3- MAKE SURE THAT PARKINBG BRAKE IS APPLIED 4- MAKE SURE THAT THE AIRCRAFT IS PROPERLY CHOCKED
HMU - WiFi Connectivity Function
TEST INHIBIT/INTERLOCK CONDITIONS OK
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Continue
Cancel
CS1_CS3_7330_004
PARKING BRAKE
Figure 35: Power Assurance Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-75
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-30 Indicating
73-30 INDICATING GENERAL DESCRIPTION The fuel indicating system consists of a fuel flow meter to monitor fuel flow and a fuel filter differential pressure sensor to provide an indication of a blocked fuel filter. The fuel flow transmitter measures metered fuel for combustion. The EEC calculates the fuel flow for display on the engine indication and crew alerting system (EICAS) page. The fuel filter differential pressure sensor measures differential pressure across the fuel filter. The fuel filter status is displayed on the FUEL synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-76
CS130_73_L2_Indicating_General.fm
73 - Engine Fuel and Control 73-30 Indicating
CLB 92.2
92.2
73.0
73.3
N1
718
722 STATUS
EGT
FUEL FUE EL
86.5 1090 124 116
FUEL FLOW METER
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
TOTAL FUEL FUEL USED
AIR IR
DOOR OR
ELECC
FLT CTR CTRLL
HYD YD
AVIONIC IONIC
INFO
CB
15300 KG 1000 KG
EICAS
FUEL FILTER DIFFERENTIAL PRESSURE SENSOR
20 °C 2900 KG
20 °C 2900 KG
ELECTRONIC ENGINE CONTROL
9500 KG
APU
ARINC 429 SYNOPTIC PAGE – FUEL
CS1_CS3_7330_002
LEGEND
Figure 36: Fuel Control Indicating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-77
CS130_73_L2_Indicating_General.fm
73 - Engine Fuel and Control 73-30 Indicating
COMPONENT LOCATION The engine fuel indicating system consists of the following components: •
Fuel flow meter
•
Fuel filter differential pressure sensor
Fuel Flow Meter The fuel flow meter is located above the integrated fuel pump and control (IFPC) on the fuel/oil manifold at the 9 o’clock position. Fuel Filter Differential Pressure Sensor The fuel filter differential pressure sensor is mounted on the forward side of the fuel/oil manifold adjacent to the main fuel filter at the 7:30 o’clock position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-78
CS130_73_L2_Indicating_ComponentLocation.fm
73 - Engine Fuel and Control 73-30 Indicating
A
A
FUEL FLOW METER
B
FUEL FILTER DIFFERENTIAL PRESSURE SENSOR
CS1_CS3_7321_011
B
Figure 37: Fuel Control Indicating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-79
CS130_73_L2_Indicating_ComponentLocation.fm
73 - Engine Fuel and Control 73-30 Indicating
CONTROLS AND INDICATIONS The fuel flow information is displayed on the EICAS page. The fuel filter bypass status is indicated on the FUEL synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-80
CS130_73_L2_Indicating_ControlsIndications.fm
73 - Engine Fuel and Control 73-30 Indicating
CLB
STATUS
92.2
92.2
73.0
73.3
FUEL TOTAL FUEL FUEL USED
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
FUEL STATUS
9500 KG 3000 KG
Symbol
Condition Normal
N1
718
722 Fuel Filter Impending Bypass
EGT
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
5000 KG
Fuel Filter Bypass APU
BYPASS
SYNOPTIC PAGE – FUEL
CS1_CS3_7322_003
EICAS
23 °C 2000 KG
23 °C 2500 KG
Figure 38: Fuel Control Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-81
CS130_73_L2_Indicating_ControlsIndications.fm
73 - Engine Fuel and Control 73-30 Indicating
DETAILED DESCRIPTION The fuel flow meter sends a signal to channel A of the EEC where it is used to calculate the metered fuel flow. The signal is supplied to channel B internally via a CAN BUS. The fuel filter differential pressure sensor provides an output to each channel of the EEC. The fuel filter differential pressure sensor measures differential pressure across the fuel filter. A fuel filter impending bypass is detected and displayed when the differential pressure exceeds 22 psi on either engine. An L(R) ENG FUEL FILTER advisory message along with an L (R) ENG FUEL FILTER - IMPENDING BYPASS info message. A fuel filter bypass occurs at 25 psid. This is indicated by the L(R) ENG FUEL FILTER advisory message and an L (R) ENG FUEL FILTER BYPASS info message. If both fuel filters are bypassed, fuel contamination can be suspected and an L-R ENG FUEL FILTER caution message is displayed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-82
CS130_73_L3_Indicating_DetailedDescription.fm
73 - Engine Fuel and Control 73-30 Indicating
CLB
L ENG FUEL FILTER
92.2
92.2
73.0
ENG FUEL FILTER - IMPENDING BYPASS
73.3
N1
SYNOPTIC PAGE - INFO 718
722
EGT
86.5 1090 124 116
FUEL FLOW METER
86.6 1110 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
10905
TOTAL FUEL (KG) 2735
FUEL FILTER DIFFERENTIAL PRESSURE SENSOR
1400 5.7 ELEV 570
2735 0
CAB ALT
RATE
ó3
CREW OXY 1850
LDG TEMP (°C)
EEC
5435
24
TRIM NU STAB
HI
23
23
ND
4.2 NL RUDDER NR
INFO
Channel A
Channel A EICAS
Channel B
Channel B
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
15300 KG 1000 KG
To Fuel/Oil Manifold
20 °C 2900 KG
20 °C 2900 KG
FUEL FILTER
LEGEND EEC
APU
Electronic Engine Control ARINC 429 CAN BUS SYNOPTIC PAGE – FUEL
CS1_CS3_7330_001
9500 KG
Figure 39: Fuel Control Indicating Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-83
CS130_73_L3_Indicating_DetailedDescription.fm
73 - Engine Fuel and Control 73-30 Indicating
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine fuel indication.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-84
CS130_73_L2_Indicating_MonitoringTests.fm
73 - Engine Fuel and Control 73-30 Indicating
CAS MESSAGES Table 7: INFO Messages
Table 5: CAUTION Messages MESSAGE
MESSAGE
LOGIC
L - R ENG FUEL FILTER
Both engine fuel filters are bypassed (clogged filters), or impending, i.e. likely fuel contamination.
Table 6: ADVISORY Messages MESSAGE
LOGIC
L ENG FUEL FILTER
Left engine fuel filter is impending bypass or is bypassed.
R ENG FUEL FILTER
Right engine fuel filter is impending bypass or is bypassed.
L ENGINE FAULT
Loss of redundant or non-critical function for the left engine.
R ENGINE FAULT
Loss of redundant or non-critical function for the right engine.
L FUEL FLOW DEGRADED
FADEC provides synthesized fuel flow instead of measured fuel flow. Displayed L fuel flow accuracy is degraded.
R FUEL FLOW DEGRADED
FADEC provides synthesized fuel flow instead of measured fuel flow. Displayed R fuel flow accuracy is degraded.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
73 L (R) ENGINE FAULT - FUEL FILTER PRESS SNSR INOP
FADEC logic has determined that the fuel filter delta P sensor signal from channel A and/or B has failed.
73 L (R) ENG FUEL FILTER IMPENDING BYPASS
INFO message will be set by the left (right) engine when the fuel filter impending bypass is detected but there is no actual bypass.
73 L (R) ENG FUEL FILTER BYPASS
INFO message will be set by the left (right) engine when the fuel filter actual bypass is detected.
TECHNICAL TRAINING MANUAL
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73-85
CS130_73_L2_Indicating_MonitoringTests.fm
73 - Engine Fuel and Control 73-30 Indicating
Page Intentionally Left Blank
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73-86
CS130_73_ILB.fm
ATA 74-80 - Ignition and Starting
BD-500-1A10 BD-500-1A11
74 - Ignition and Starting
Table of Contents 74-11 Ignition ....................................................................74-80-2 General Description .....................................................74-80-2 Component Location ...................................................74-80-4 Ignition Exciter .......................................................74-80-4 Igniter Plugs ............................................................74-80-4 Ignition Cables ........................................................74-80-4 Component Information ...............................................74-80-6 Ignition Exciter .......................................................74-80-6 Ignition Cables .......................................................74-80-8 Igniter Plugs ............................................................74-80-8 Controls and Indications ............................................74-80-10 Engine Panel.........................................................74-80-10 Throttle Quadrant Assembly .................................74-80-10 Indications.............................................................74-80-12 Detailed Description ..................................................74-80-14 Normal Ignition for Automatic Start .......................74-80-14 Continuous Ignition ...............................................74-80-14 Monitoring and Tests .................................................74-80-18 CAS Messages .....................................................74-80-19 Practical Aspects........................................................74-80-20 Ignition System Test ............................................74-80-20 80-11 Starting .................................................................74-80-22
Component Information ............................................ 74-80-26 Air Turbine Starter................................................ 74-80-26 Starter Air Valve ................................................... 74-80-28 Controls and Indications ........................................... 74-80-30 Throttle Quadrant Assembly ................................ 74-80-30 Engine Panel........................................................ 74-80-30 Indications ............................................................ 74-80-32 Operation .................................................................. 74-80-34 Rotor Bow ............................................................ 74-80-34 Automatic Start..................................................... 74-80-36 Manual Start......................................................... 74-80-38 Engine Motoring ................................................... 74-80-38 Detailed Description ................................................. 74-80-40 Automatic Starting................................................ 74-80-40 Starter Duty Time ................................................. 74-80-40 Starter Re-Engagement ....................................... 74-80-40 Abnormal Starts ................................................... 74-80-42 Monitoring and Tests ............................................... 74-80-46 CAS Messages .................................................... 74-80-47 Practical Aspects ....................................................... 74-80-48 Starter Oil Servicing ............................................ 74-80-48 Starter Chip Collector........................................... 74-80-48
General Description ...................................................74-80-22 Automatic Start .....................................................74-80-22 Manual Start..........................................................74-80-22 Rotor Bow .............................................................74-80-22 Component Location .................................................74-80-24 Air Turbine Starter.................................................74-80-24 Starter Air Valve....................................................74-80-24 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-i
CS130_M5_74-80_IgnitionStartingTOC.fm
74 - Ignition and Starting
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7: Figure 8: Figure 9: Figure 10: Figure 11:
Ignition System.................................................74-80-3 Ignition System Component Location...............74-80-5 Ignition Exciter..................................................74-80-7 Ignition Cables and Igniter Plugs......................74-80-9 Ignition System Controls ................................74-80-11 Ignition System Indications.............................74-80-13 Ignition System Schematic .............................74-80-15 Ignition and Starting Modes............................74-80-17 Ignition System Test.......................................74-80-21 Engine Starting System..................................74-80-23 Engine Starting System Component Location ......................................74-80-25 Figure 12: Air Turbine Starter ..........................................74-80-27 Figure 13: Starter Air Valve .............................................74-80-29 Figure 14: Engine Start Controls .....................................74-80-31 Figure 15: Engine Starting Indications.............................74-80-33 Figure 16: Rotor Bow Altitude/Outside Air Temperature Graph ..................................74-80-35 Figure 17: Automatic Start...............................................74-80-37 Figure 18: Manual Start ...................................................74-80-39 Figure 19: Automatic and Manual Starting Schematic ....74-80-41 Figure 20: Abnormal Starts..............................................74-80-43 Figure 21: Fuel Depulsing ...............................................74-80-45 Figure 22: Starter Oil Servicing and Starter Chip Collector .....................................74-80-49
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74-ii
CS130_M5_74-80_IgnitionStartingLOF.fm
74 - Ignition and Starting
Page Intentionally Left Blank
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74-iii
CS130_M5_74-80_IgnitionStartingLOF.fm
IGNITION and STARTING - CHAPTER BREAKDOWN IGNITION AND STARTING CHAPTER BREAKDOWN
Ignition System
1
Starting System
2
74-80 - Ignition and Starting 74-11 Ignition
74-11 IGNITION GENERAL DESCRIPTION The engine ignition system consists of two electrically and physically independent circuits located in a single ignition exciter. Each exciter circuit is an AC-powered capacitor discharge type that supplies high voltage to an igniter plug through a shielded, high tension cable.
Automatic selection of continuous dual-ignition is provided by the EEC during critical or adverse conditions. The EEC also includes an automatic relight system, which energizes both igniters within 2 seconds when an engine flameout is detected.
Each exciter circuit is powered from the aircraft. The left engine ignition power is supplied by DC ESS BUS 1 and DC ESS BUS 3. Right engine ignition power is supplied by DC ESS BUS 2 and 3. The electronic engine control (EEC) provides the control of the ignition.
The EEC automatically selects a dual-igniter continuous ignition if any of the following conditions are true:
The ignition system is used during engine start or engine relight. The ignition system is also used any time ambient conditions require a continuous ignition to prevent the risk of flameout. During a normal automatic engine start, the EEC uses a single igniter plug to start the engine. The igniter plug selected alternates on each normal automatic engine start. The EEC commands the igniter plugs on or turns the igniter plugs off based on N2 speed.
•
Engine flameout is detected in-flight or during takeoff
•
Surge is detected in-flight or during takeoff
•
In-flight start
The automatic selection of continuous ignition occurs even if manual continuous ignition is selected.
The ignition system can also be operated in a dual-igniter low power continuous ignition mode. The continuous ignition mode operates automatically under the control of the EEC, or manually when selected from the CONT IGNITION PBA on the ENGINE panel. When the CONT IGNITION PBA is selected before engine start, the EEC enters the manual start mode and both igniter plugs are selected for engine start. In the manual start mode, the ignition is commanded on when the THROTTLE QUADRANT ASSEMBLY L(R) ENG switch is selected to the ON position, and turned off when the engine self-sustaining speed is reached. When continuous ignition is selected and the engine is already running, the igniter plugs only fire when the engine is operating at low power settings. This increases the life of the igniter plugs. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-2
CS130_74-80_L2_Ignition_General.fm
74-80 - Ignition and Starting 74-11 Ignition
MAX
MAX
IDLE
ID IDLE
L ENG ON
OFF
THROTTLE QUADRANT ASSEMBLY ENGINE CONT IGNITION
IGNITER PLUG A
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
DMC 1 and DMC 2
IGNITION EXCITER
EEC
IGNITER PLUG B ENGINE PANEL
DC ESS BUS 3
NOTE
DMC EEC
LEGEND
Left engine shown. Right engine similar.
Data Concentrator Unit Module Cabinet Electronic Engine Control
Right engine ignition exciter power DC ESS BUS 2 and DC ESS BUS 3
CS1_CS3_7400_002
DC ESS BUS 1
Figure 1: Ignition System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-3
CS130_74-80_L2_Ignition_General.fm
74-80 - Ignition and Starting 74-11 Ignition
COMPONENT LOCATION The ignition system includes the following components: •
Ignition exciter
•
Ignition cables
•
Igniter plugs
IGNITION EXCITER The ignition exciter is located on the left side of the engine fan case at the 8 o’clock position. IGNITER PLUGS The igniter plugs are installed in the diffuser case. Igniter A is located at the 4 o’clock position, and igniter B is located at the 5 o’clock position. IGNITION CABLES The ignition cables extend under the engine from the ignition exciter to the igniter plugs.
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74-80-4
CS130_74-80_L2_Ignition_ComponentLocation.fm
74-80 - Ignition and Starting 74-11 Ignition
Igniter Plugs
Ignition Cables
CS1_CS3_7400_019
Ignition Exciter
Figure 2: Ignition System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-5
CS130_74-80_L2_Ignition_ComponentLocation.fm
74-80 - Ignition and Starting 74-11 Ignition
COMPONENT INFORMATION IGNITION EXCITER The dual-channel ignition exciter supplies 5 kV to the igniter plugs. The ignition exciter has a duty cycle that varies between 1 spark per second to 3 sparks per second. The ignition exciter is cooled by the fan compartment ventilation air.
WARNING MAKE SURE THE IGNITION SYSTEM HAS NOT BEEN OPERATED FOR AT LEAST 5 MINUTES PRIOR TO REMOVING THE IGNITER PLUGS OR THEIR CABLES. IGNITION VOLTAGE CAN BE VERY HIGH AND COULD CAUSE INJURY.
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74-80-6
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
Ignition Cable B
Ignition Cable A
CS1_CS3_7400_005
Ignition Exciter
Figure 3: Ignition Exciter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-7
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
IGNITION CABLES Ignition cables are interchangeable with flexible braided steel and ceramic insulated terminals at the plug end. IGNITER PLUGS The igniter plugs extend through the diffuser case into the combustion chamber. Classified spacers are installed under a mounting boss and are used to control the immersion depth of the igniter plug. The spacer and boss are installed on the diffuser case assembly and are not removed during igniter plug replacement.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-8
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
Igniter Plug System A
A
Igniter Plug
Mounting Boss
Classified Spacer Igniter Plug System B
Ignition Cables IGNITER PLUG
CS1_CS3_7400_006
A
Figure 4: Ignition Cables and Igniter Plugs Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-9
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
CONTROLS AND INDICATIONS ENGINE PANEL The CONT IGNITION PBA is used to select manual ignition. The manual selection is indicated by the white ON legend on the pushbutton annunciator (PBA). THROTTLE QUADRANT ASSEMBLY The L(R) ENG switch signals the EEC to command the ignition ON during a manual engine start when they are selected to the ON position.
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74-80-10
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition and Starting 74-11 Ignition
ENGINE
ID D IDLE
IDLE
CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
L ENG
RE ENG NG
ENGINE PANEL
ON
OFF
OFF CS1_CS3_7400_007
THROTTLE QUADRANT ASSEMBLY
Figure 5: Ignition System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-11
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition and Starting 74-11 Ignition
INDICATIONS An IGN icon is displayed below the exhaust gas temperature (EGT) gauges when ignition is selected. The IGN icon is green when the ignition system is commanded and the igniters are firing. The IGN icon is white when continuous ignition is selected ON, but inhibited.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-12
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition and Starting 74-11 Ignition
TO 90.8
90.8
0.0
0.0
N1
44
16
EGT
SYMBOL
COLOR
DESCRIPTION
IGN
Green
Ignition active (Firing).
IGN
White
Ignition selected but inhibited due to design constraints.
IGN N2 FF (KPH) OIL TEMP OIL PRESS FAN VIB
0.0 0 16 0 ---
CS1_CS3_7400_020
28.0 90 17 47 ---
EICAS PAGE
Figure 6: Ignition System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-13
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition 74-11 Ignition
DETAILED DESCRIPTION When the engine is running, manual continuous ignition can be selected by cycling the CONT IGNITION PBA to NORMAL and back to ON.
NORMAL IGNITION FOR AUTOMATIC START The EEC channel A controls ignition system A, and EEC channel B controls ignition system B. A single igniter plug is used during an automatic start. During normal starts, the igniter plugs are commanded on at 20.4% to 23.5% N2. When the N2 reaches 49.3 to 53.3%, the EEC commands the igniter plugs off. A green IGN icon is shown on EICAS anytime the EEC activates the igniters. The icon is removed when the igniter plugs are not on. If the engine fails to start on the first attempt, the EEC dry motors the engine, and then selects both igniter plugs for a second start attempt.
When the engine is running and manual continuous ignition is selected, a white IGN icon is displayed on the engine indication and crew alerting system (EICAS). In this condition, the igniter plugs may not be firing unless the engine is at low power. The EEC monitors the burner pressure (Pb). If Pb is greater than 150 psi, the ignition is commanded off. Once the engine is running at approach idle or less, Pb is low and the ignition is commanded on. Automatic Continuous Ignition The EEC automatically selects dual-igniter plugs continuous ignition under any of the following conditions:
CONTINUOUS IGNITION
•
Engine flameout detected in-flight or during takeoff above 60 kt
After the engine has started, normal engine operation does not require the igniter plugs to sustain combustion. However, under certain conditions continuous ignition is available for use.
•
An engine surge is detected in air
•
In-flight start
Continuous ignition has two modes:
Quick-Relight Continuous Ignition
•
Manual
•
Automatic
The engine performs a quick-relight in-flight only when the L(R) ENG switch is cycled from ON to OFF, and back to the ON position.
Manual Continuous Ignition Manual continuous ignition is available when the CONT IGNITION PBA is pressed. An ON light illuminates on the PBA to indicate manual continuous ignition is available.
When the quick-relight condition occurs, the EEC commands continuous ignition until either the engine has been above idle for 30 seconds, or the L(R) ENG switch is placed in the OFF position.
During a manual engine start, continuous ignition is used and both igniter plugs fire once the L(R) ENG switches are selected to ON. The EEC commands the igniter plugs off when the N2 reaches 55%. A green IGN icon appears as long as the igniters are firing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-14
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition 74-11 Ignition
MAX
MAX
EPC 1 DC ESS BUS 3
ID IDLE
IDLE
EEC L ENG
7.5
ENG IGN A
IGNITION EXCITER
Channel A
R ENG ENG
ON
Ignition Command
OFF
TQA
Channel B
DMC 1 and DMC 2
EICAS
ENGINE START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
System B
IGNITER PLUG B
P3
N2
ON
CDC 1-7-1 L ENG IGN B DC ESS 1 SSPC 7.5A Logic: Always On
ENGINE PANEL
LEGEND CDC DMC EEC EPC P3 TQA
NOTE Left engine shown, right engine similar.
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Engine Control Electrical Power Center Burner Pressure Throttle Quadrant Assembly ARINC 429
CS1_CS3_7400_012
CONT IGNITION
IGNITER PLUG A
Ignition Command
IGN
Air/Gnd
System A
Figure 7: Ignition System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-15
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition 74-11 Ignition
Ignition and Starting Modes The START switch located on the ENGINE panel is in the AUTO position for normal engine starts. The L (R) ENG CRANK position is used for manual engine starts, dry motoring, and wet motoring. The CONT IGNITION PBA is used to initiate a manual engine start or select both igniters on. The L (R) ENG switches initiate an automatic start when the START switch is in AUTO. During a manual engine start, moving the L (R) ENG to ON initiates fuel and ignition. When the switch is moved to the OFF position, the engine shuts down. During an automatic engine start, only one igniter is used. The EEC alternates igniters on consecutive starts. During a manual engine start on the ground or an engine start in the air, both igniters are on. When the engine is dry or wet motored, the igniters are off. In addition to the ignition selection for engine starting, the EEC automatically selects dual-igniter ignition under any of the following conditions: •
In-flight engine start
•
Engine flameout detected in-flight or during takeoff above 60 kt
•
An engine surge is detected in air
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-16
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition 74-11 Ignition
ENGINE START
CONT IGNITION
L ENG CRANK
AUTO
R ENG CRANK
L ENG
R ENG
ON
ON
OFF FF FF
OFF FF FF
ON
CONT IGNITION NORMAL
START AUTO
L (R) ENG ON
Manual Start (Ground)
CONT IGNITION ON
START L (R) ENG CRANK
L (R) ENG ON
Auto Start (Air)
CONT IGNITION NORMAL
START AUTO
L (R) ENG ON
Dry Motoring
CONT IGNITION NORMAL
START L (R) ENG CRANK
L (R) ENG OFF
Wet Motoring
CONT IGNITION NORMAL
START L (R) ENG CRANK
L (R) ENG ON
Engine Shutdown
Alternates Igniters Every Start
Automatic Continuous Ignition • Engine Flameout on Takeoff or in Air • Engine Surge in Air
Air Turbine Starter (ATS) or Windmill Start
44
16
EGT IGN
L (R) ENG OFF
28.0
N2
0.0
EICAS PAGE WITH ENGINE IGN ICONS
CS1_CS3_7400_021
Auto Start (Ground)
Figure 8: Ignition and Starting Modes Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-17
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition and Starting 74-11 Ignition
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the starting and ignition system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-18
CS130_74-80_L2_Ignition_MonitoringTests.fm
74-80 - Ignition and Starting 74-11 Ignition
CAS MESSAGES Table 1: STATUS Message MESSAGE ENG CONT IGNITION ON
LOGIC At least one EEC receives a request from the flight deck switch to have continuous ignition.
Table 2: INFO Message MESSAGE 74 L(R) ENGINE FAULT - IGN REDUND LOSS
LOGIC An ignition circuit fault, or failure to light using an igniter, or a low side voltage fault has been detected by the left/right full authority digital engine control (FADEC) system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-19
CS130_74-80_L2_Ignition_MonitoringTests.fm
74-80 - Ignition and Starting 74-11 Ignition
PRACTICAL ASPECTS IGNITION SYSTEM TEST The ignition system test procedure checks for satisfactory operation of the ignition system. The following functions or units are verified with this test: •
EEC outputs to exciter
•
Exciter
•
Ignition cables (two per engine)
•
Igniter plugs (two per engine)
The engine igniters are wired to one channel each, with igniter 1 going to EEC channel A, and igniter 2 going to EEC channel B. As a consequence, the ignition system test can only verify the functionality of the igniter corresponding to the channel in control during the test. To test both igniters, the maintainer must run the test again on the alternate EEC channel. This test allows the user to audibly verify if the igniters provide a spark when commanded to do so.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-20
CS130_74-80_L2_Ignition_PracticalAspects.fm
74-80 - Ignition and Starting 74-11 Ignition
MAINT MENU EEC1-A-IGNITION TEST
1/4
*** WARNING *** IGNITERS WILL SPARK DURING THIS TEST A DRY CRANK MUST BE PERFORMED PRIOR TO RUNNING THIS TEST TO ENSURE NO ACCIDENTAL FUEL LIGHT OFF
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- PERFORM DRY CRANK AMP XX-XX-XX PRIOR TO THE TEST 3- FADEC IS POWERED
* NOTE * THERE IS A 5 MINUTE TIMER ON THE OPERATION OF THE IGNITER UPON PRESSING THE START BUTTON
PRESS TO START TEST
Write Maintenance Reports
Cancel
CS1_CS3_7321_023
PRESS THE START BUTTON BELOW TO INITIATE TEST
Figure 9: Ignition System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-21
CS130_74-80_L2_Ignition_PracticalAspects.fm
74-80 - Ignition and Starting 80-11 Starting
80-11 STARTING GENERAL DESCRIPTION
MANUAL START
The engine starting system consists of an air turbine starter (ATS) and a starter air valve (SAV). The ATS uses bleed air to drive the high-pressure (HP) rotor to a high enough speed to ensure a satisfactory start. The bleed air is provided by the auxiliary power unit (APU), an external ground air cart, or the other operating engine. The SAV controls the flow of bleed air to the ATS. When the engine has reached a predetermined speed, the SAV closes, shutting off the bleed air to the ATS. As the engine continues to accelerate, the starter clutch automatically disengages the starter. The ATS can also be used to start the engine in flight if a windmill start can not be accomplished. Engine starting can be accomplished automatically or manually. AUTOMATIC START The automatic start mode provides the electronic engine control (EEC) with full control to sequence the SAV, ignition exciter and the fuel metering valve located in the integrated fuel pump and control (IFPC). The automatic start is initiated when the engine is not running, the START switch is in AUTO, and the L(R) ENG switch is set to ON. When the start command is received from the L(R) ENG switch, the EEC commands the SAV, ignition exciter, and the fuel-on function of the IFPC. Once the engine reaches the starter cutout speed, the EEC commands the SAV closed, and the ignition exciter off. Switching the L(R) ENG switch OFF during an automatic start shuts off the fuel, closes the SAV, and turns the ignition system off. The EEC monitors the automatic start and has the capability to abort the startup to the self-sustaining engine speed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Manual engine starting is available on the ground only. The manual start mode limits the authority of the EEC so that the starter, ignition, and fuel are controlled manually. This provides the ability to dry motor or wet motor the engine. The manual start sequence is available when the CONT IGNITION PBA is selected in the ON position. The start is initiated when the START switch is selected to the L (R) ENG CRANK position. The EEC commands the SAV open and the ATS drives the HP rotor. Once the engine is rotating, fuel and ignition exciter are commanded ON by selecting the L(R) ENG switch to ON. Once the engine reaches the starter cutout speed, the EEC commands the SAV closed, and the ignition exciter off. The manual start sequence can be terminated by selecting the START switch to the AUTO position if the L(R) ENG switch has not been selected ON, or by selecting the L(R) ENG switch to OFF. The EEC only provides fault indications for the engine indication and crew alerting system (EICAS) during manual operation. The EEC start abort function is not available. Monitoring of the start parameters is required. The start is terminated by selecting the L(R) ENG switch to OFF. ROTOR BOW To minimize the effects of a bowed rotor, a motor-to-start dry crank is incorporated into every engine start sequence. The starter air valve is opened and modulated to maintain an N2 speed between 7% and 13% for a minimum of 30 seconds before continuing with the engine start.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-22
CS130_74-80_L2_Starting_General.fm
74-80 - Ignition and Starting 80-11 Starting
DC ESS BUS 1 DC ESS BUS 3 ENGINE CONT IGNITION
L ENG
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
ON
OFF
ENGINE PANEL
THROTTLE QUADRANT ASSEMBLY
IGNITION EXCITER
DMC 1 AND DMC 2
EEC
IFPC
Air Data
DMC EEC IFPC T/M
Bleed Air Data Concentrator Unit Module Cabinet Electronic Engine Control Integrated Fuel Pump and Control ARINC 429 Torque Motor
T/M
AIR TURBINE STARTER
Bleed Air System
MAIN GEARBOX
STARTER AIR VALVE NOTE Left engine shown, right engine similar.
CS1_CS3_8000_004
LEGEND
N2 Speed
Figure 10: Engine Starting System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-23
CS130_74-80_L2_Starting_General.fm
74-80 - Ignition and Starting 80-11 Starting
COMPONENT LOCATION AIR TURBINE STARTER The air turbine starter (ATS) is mounted to the main gearbox (MGB), located on the left side of the engine below the air oil cooler. STARTER AIR VALVE The starter air valve (SAV) is mounted to the MGB, located on the left side of the engine behind the air cooler.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-24
CS130_74-80_L2_Starting_ComponentLocation.fm
74-80 - Ignition and Starting 80-11 Starting
Starter Air Valve
CS1_CS3_8000_005 00 005
Air Turbine Starter
Figure 11: Engine Starting System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-25
CS130_74-80_L2_Starting_ComponentLocation.fm
74-80 - Ignition and Starting 80-11 Starting
COMPONENT INFORMATION AIR TURBINE STARTER The starter converts power from the bleed air system to rotational energy using a turbine and a reduction gearbox. The gearbox drives the output shaft through a ratchet and pawl clutch. At low engine speeds, the pawl springs force the pawls to engage with the ratchet on the output shaft. When the engine speed exceeds that of the output hub, the pawls act as fly weights and disengages from the drive shaft. To prevent damage to the air turbine stater (ATS) or the engine during start, the output shaft has a shear section between it and the main gearbox (MGB). This is designed to fail within a specific torque range. Starter speed is sent to the EEC by a dual-channel speed sensor located on the starter housing. The speed sensor is a line replaceable unit (LRU) that provides a starter speed signal to each channel of the engine electronic control (EEC). The EEC uses the starter speed to determine start valve position, and starter engagement and disengagement. The starter has a sight glass to verify the oil level, a fill port for servicing the oil, and a starter chip collector. The starter chip collector is to prevent any debris from circulating through the ATS oil system. A locator pin on the ATS gearbox interface is used to ensure the ATS is properly clocked during installation. The ATS is cooled using fan air from the inlet of the air/oil cooler. The air flows through a cooling ring around the ATS forward flange.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-26
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
FULL
Locator Pin ADD
Output Shaft
A
SIGHT GLASS
A
Dual-Channel Speed Sensor
Starter Sight Glass Chip Collector
Fill Port
CS1_CS3_8000_006
Cooling Ring
Figure 12: Air Turbine Starter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-27
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
STARTER AIR VALVE
Manual Override Start Procedure
The starter air valve (SAV) is a pneumatically-actuated butterfly valve. A pressure regulator maintains an output pressure of 30 psi. The valve is spring-loaded to the closed position. The SAV is EEC controlled using a dual-coil torque motor. During start, the valve opening is confirmed through starter speed (Ns) and the N2.
The following procedure is used during a manual override start:
A filter port is accessible once the shroud is removed to access a line replaceable filter.
4. Ensure the SAV is in the closed position. Rotate the square drive clockwise to confirm.
A cooling shroud surrounds the SAV. A hose provides air from the air/oil cooler to the shroud to protect the valve from the heat of the engine core.
5. When directed, open the SAV by turning the manual override counterclockwise until the stop is contacted.
If the SAV fails to operate, a manual override is provided to open the valve. An SAV manual override access hole on the cowls permits a 3/8 in square drive manual override to be inserted into the SAV for manual opening.
6. Hold the valve in the open position.
1. Establish communication with the flight deck. 2. Ensure bleed air is available to the SAV. 3. Insert the 3/8 in square drive into the SAV manual access hole.
7. When directed, allow the valve to return to the closed position. The valve is spring-loaded to the closed position. Rotate the square drive clockwise to close if required. 8. Remove the manual override drive. 9. Move clear of the aircraft, avoiding the engine hazard areas. 10. Contact the flight deck to indicate clear of the aircraft.
NOTE This is not an approved procedure at this time due to the incorporation of the motor-to-start logic for rotor bow prevention.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-28
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
Cooling Shroud (4x)
Starter Air Valve Manual Override Starter Air Valve (SAV) Override Access Hole
Torque Motor
CS1 CS3 8000 008 CS1_CS3_8000_008
Filter Port
Figure 13: Starter Air Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-29
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
CONTROLS AND INDICATIONS THROTTLE QUADRANT ASSEMBLY The L(R) ENG switches are used to initiate an automatic start, or command the fuel and ignition during a manual start. The L(R) ENG switches are also used to shut down the engine. ENGINE PANEL The engine panel has a three-position START switch marked AUTO and L(R) ENG CRANK. The switch is spring-loaded to the AUTO position. When a start is initiated with the switch in the AUTO position, an automatic start is performed. The L(R) ENG CRANK positions are used for manual starts and wet or dry motoring. The CONT IGNITION PBA is used to select manual ignition. It also selects the manual start mode for the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-30
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
ENGINE
ID D IDLE
IDLE
CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
L ENG
ENGINE PANEL
ON
OFF
OFF CS1_CS3_8000_010
THROTTLE QUADRANT ASSEMBLY
Figure 14: Engine Start Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-31
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
INDICATIONS While the EEC controls the start sequence, the engine indication and crew alerting system (EICAS) page is monitored during the start sequence. Indications that are monitored during engine start include: •
N1
•
Exhaust gas temperature (EGT)
•
Ignition
•
N2
•
Fuel flow
•
Oil pressure
The following indications are displayed above the N1 gauges, depending upon start mode: •
START - Engine is in the start sequence
•
RELIGHT - Indicates an automatic relight
In the case of an automatic restart in flight, there are the following indications depending on the type of restart: •
WINDMILL - Altitude from 30,000 ft to sea level, airspeed above 250 KIAS
•
ATS - Indicates an air turbine starter start mode. There are two different types of ATS starts -
Starter assisted APU - Altitude from 23,000 ft to sea level
-
Engine cross-bleed - Altitude from 30,000 ft to sea level, airspeed less than 250 KIAS
The EEC determines what type of in-flight restart to use depending on the altitude and airspeed of the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-32
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
TO
START 90.8
90.8
0.0
0.0
N1 START ICONS
44
16
RELIGHT ATS
EGT IGN
28.0 90 17 47 ---
START
WINDMILL N2
FF (KPH) OIL TEMP OIL PRESS FAN VIB
0.0 0 16 0 ---
EICAS PAGE
Description Engine in start sequence. Automatic relight. ATS (Air Turbine Start). Windmill.
CS1_CS3_7400_008
Symbol
Figure 15: Engine Starting Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-33
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
OPERATION ROTOR BOW The motor-to-start time for rotor bow prevention is based on an altitude/outside air temperature (OAT) graph. Within the altitude/OAT envelope, the motor to start time is 30 seconds. As the distance outside the altitude/OAT envelope increases, the motor-to-start time increases. An ENG START DELAY advisory message is displayed anytime an engine start is initiated outside of the envelope. The message clears when the motor to start phase is complete.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-34
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
ALT
14000 ft
30 Second Motor-To-Start Time
5000 ft
ENG START DELAY
-2000 ft
OAT -17°C
-5°C
+43°C
CS1_CS3_7400_022
ENG START DELAY
Figure 16: Rotor Bow Altitude/Outside Air Temperature Graph Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-35
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
AUTOMATIC START When the L(R) ENG switch is moved to the ON position, the electronic engine control (EEC) begins the start sequence: •
Starter air valve (SAV) is opened
•
Oil pressure increases
•
START icon displayed
•
ENG START DELAY advisory message on if motor-to-start time > 30 seconds
•
Engine dry cranks between 7%-13% N2 for a minimum of 30 seconds
•
ENG START DELAY message off
•
At 20.4% to 23.5% N2, the EEC commands ignition using one of the igniter plugs
•
IGN icon on
•
Commands the fuel metering valve in the integrated fuel pump and control (IFPC) to supply a fuel flow of 91 kg/hr (200 lb/hr)
•
Within 20 seconds of fuel flow: -
EGT increases
-
N1 increases before 45% N2
•
At 49.3 to 53.3% N2, the EEC commands the starter air valve closed and shuts off the ignition
•
IGN icon off
•
At engine idle, 62% N2, fuel flow reaches 227 kg/hr (500 lb/hr)
•
START icon off
For automatic starts in flight, the EEC selects both igniters on.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-36
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
100
80 ENG START DELAY Advisory Message OFF
N2 (%) 40
20
SAV Modulates to Maintain 7-13% N2
0
49.3% to 53.3% N2 45% N2
• EGT rise • Fuel flow increase • N1 increase after 38% N2
30 seconds minimum
Engine IDLE FF: 227 kg/hr START Icon OFF
62% N2
20.4 to 23.5% N2 EEC commands: • Fuel ON • Ignition ON • IGN Icon ON
60
49.3% to 53.3% N2 EEC commands: • SAV closed • Ignition OFF • IGN Icon OFF
15 seconds TIME →
EEC EGT FF kg/hr SAV
Electronic Engine Control Exhaust Gas Temperature Fuel Flow Kilograms per Hour Starter Air Valve
• L(R) ENG Switch to ON • Starter Air Valve Open • START Icon On • ENG START DELAY Advisory Message on if motor-to-start time > 30 seconds • Oil Pressure Increases
L ENG ON
ENGINE CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
ON
OFF
THROTTLE QUADRANT ASSEMBLY
ENGINE PANEL
PILOT EVENT
NOTE Three starter crank cycles for a total of four minutes cranking time maximum.
CS1_CS3_7400_010
LEGEND
Figure 17: Automatic Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-37
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
MANUAL START In a manual start, the CONT IGNITION PBA is selected to ON. The START switch is held in the L(R) ENG CRANK position, and the following occurs; •
Starter air valve opens
•
ENG START DELAY advisory message on if motor-to-start time > 30 seconds
•
Engine dry cranks between 7%-13% N2 for a minimum of 30 seconds
•
ENGINE START DELAY message off
•
At 18.5% to engine maximum motoring speed of 24% N2, the L(R) ENG switch is moved to RUN
•
•
When the engine idle speed is reached the CONT IGNITION PBA is selected NORMAL. ENGINE MOTORING When on ground, the engine can be motored for maintenance purposes. The engine can run continually for 4 minutes without stopping. A wait time of 30 minutes is required before starting again. There are two motoring sequences, dry and wet motoring. Dry Motoring The EEC initiates a dry motoring when bleed air is available and the following controls are set: •
START switch in L(R) ENG CRANK
-
START icon on
•
CONT IGNITION PBA in normal
-
The EEC commands the ignition using both igniter plugs
•
L(R) ENG switch OFF
-
IGN icon on
-
Commands the fuel metering valve in the integrated fuel pump and control (IFPC) to open and supply a fuel flow of 91 kg/hr (200 lb/hr)
The motoring can be interrupted any time by releasing the START switch to AUTO.
The EEC initiates a wet motoring when bleed air is available and the following controls are set:
Within 20 seconds of fuel flow: -
EGT increases
-
N1 increases before 45% N2
Wet Motoring
•
Pull the applicable ignition circuit breakers or open the ignition solid-state power controllers (SSPCs)
At 49.3 to 53.3% N2, the START switch is released to the AUTO position then the EEC commands the SAV closed and shuts off the ignition
•
START switch in L(R) ENG CRANK
•
CONT IGNITION PBA in normal
-
•
L(R) ENG switch to RUN
IGN icon OFF
•
At engine idle, 62% N2, fuel flow reaches 227 kg/hr (500 lb/hr)
•
Start icon off
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Positioning the L(R) ENG switch to the OFF position shuts off fuel and initiates a dry crank. The motoring can be interrupted any time by releasing the engine START switch to AUTO.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-38
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
100
80
60
18.5 to 24% N2 • L(R) ENG Switch to RUN • START Icon On EEC commands: • Fuel ON • Ignition ON • IGN Icon ON
ENG START DELAY Advisory Message OFF
N2 (%)
Engine IDLE FF: 227 kg/hr START Icon OFF
62% N2 49.3 to 53.3% N2 45% N2
40 Oil Pressure Increasing
20
0
SAV Modulates to Maintain 7-13% N2
• EGT rise • Fuel flow increase • N1 increase after 38% N2
30 seconds minimum
49.3 to 53.3% N2 START Switch to AUTO EEC commands: • SAV closed • Ignition OFF • IGN Icon OFF
15 seconds TIME →
EEC EGT FF kg/hr SAV
L ENG ON
ENGINE CONT IGNITION
NOTE
START L ENG CRANK
AUTO
R ENG CRANK
ON
OFF F
THROTTLE QUADRANT ASSEMBLY
ENGINE PANEL
PILOT EVENT
Three starter crank cycles for a total of four minutes cranking time maximum. Four minutes maximum continuous cranking.
CS1_CS3_7400_017
• CONT IGNITION PBA ON • START Switch to L(R) ENG CRANK Electronic Engine Control • Starter Air Valve Open Exhaust Gas Temperature • ENG START DELAY Advisory Fuel Flow Message on if motor-to-start time Kilograms per Hour > 30 seconds Starter Air Valve
LEGEND
Figure 18: Manual Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-39
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
DETAILED DESCRIPTION
not reached and the START icon disappears. The message clears when the L(R) ENG switch is selected to OFF.
AUTOMATIC STARTING
The EEC start abort is inhibited above 49.3% N2 to 53.3% N2 on the ground and at all conditions in-flight. Following a ground start abort, an automatic 30 seconds engine motoring period is provided to clear fuel vapor and to cool the engine. An automatic start performs one restart attempt on the ground.
During engine start, the EEC modulates the SAV to maintain the N2 speed between 7% and 13% for a minimum of 30 seconds to prevent rotor bow. An ENG START DELAY advisory message is displayed if the motor to start time exceeds 30 seconds. During a normal start, the starter air valve (SAV) and ignition exciter are automatically turned off by the electronic engine control (EEC) between 49.3% and 53.3% N2 speed, depending on ambient conditions. A normal engine start is indicated by the START icon above the N1 indicator. Starter assist is commanded by the EEC for in-flight starts at low Mach numbers where windmilling conditions are insufficient for engine starting. The EEC selects the starter assistance mode if altitude or Mach number are not available. During an in-flight start, the EEC depressurizes the engine-driven hydraulic pump (EDP) to unload the engine and improve starting performance. A windmill start is carried out if the altitude is between 30,000 ft and sea level with airspeed above 250 KIAS. If the aircraft is not in the windmill start envelope, the starter can be used to assist the start. If the altitude is between 30,000 ft and 23,000 ft, the engine can be restarted using bleed air from the opposite engine. If the altitude and speed are below 23,000 ft and less than 250 kt, the auxiliary power unit (APU) can be used for starting. On the ground, upon detection of a no light, hot or hung start, loss of exhaust gas temperature (EGT) or a locked low-pressure rotor, the EEC automatically shuts off fuel, ignition, and starter air and provides the appropriate fault indication to the flight deck. An automatic start abort is indicated by an L(R) ENG START ABORT caution message on the engine indication and crew alerting system (EICAS). The message indicates that the engine start procedure has aborted. Idle speed was
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
STARTER DUTY TIME A STARTER TIME EXCEEDED caution message is displayed during auto or manual starts when the starter operating time reaches 4 minutes, or if the starter is operated above 44% N2 for 30 seconds. A 30 minute cooling time is required after three start attempts or 4 minutes of continuous cranking. The EEC monitors the starter operating time including during both wet and dry cranking. During an automatic start on the ground, the start attempt is discontinued and the EEC closes the starter air valve. When the message occurs, the L(R) ENG switch must be selected OFF to reset the EEC timers. For a manual start or an in-flight start, the message remains displayed until the start is terminated, or the engine reaches idle. STARTER RE-ENGAGEMENT The L(R) ENG switch must be selected OFF before a second attempt can be made following an aborted manual start attempt or an engine shutdown. If the starter engagement is interrupted during a manual start, the starter can be re-engaged below the starter cut-out speed. During manual or automatic restarts, the EEC prevents starter re-engagement until the starter rotational speed is synchronized with engine N2 speed. The starter re-engagement speed is 20% N2. The EEC compares N2 with the starter speed signal. If the starter speed is unavailable, the start valve must be closed for 1 minute, then the starter can be re-engaged at 20% N2 on the ground, or in the air.
TECHNICAL TRAINING MANUAL
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74-80-40
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
ENGINE CONT IGNITION
L ENG
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
ON
OFF
EDP
THROTTLE QUADRANT ASSEMBLY
ENGINE PANEL
EEC
Channel A
IGN EICAS
DMC 1 and DMC 2
IFPC
Channel B IGNITION EXCITER
ATS Air Data • Altitude • MACH
Starter Speed
ATS DMC EDP EEC IFPC
Air Turbine Starter Data Concentrator Unit Module Cabinet Engine-Driven Hydraulic Pump Electronic Engine Control Integrated Fuel Pump And Control ARINC 429
STARTER AIR VALVE
CS1_CS3_8011_001
LEGEND
Figure 19: Automatic and Manual Starting Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-41
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
ABNORMAL STARTS Normal Engine Start
Automatic Start Abort
If the oil pressure is less than 70 psi after 30 seconds at the usual idle speed, shut down the engine using the normal shutdown procedure.
The EEC aborts an automatic start if the engine encounters one or more of the following events during the start, and are not recoverable using the fuel depulse or automatic restart logic.
The engine should be operated at idle power for a minimum of 5 minutes before shutdown. If necessary, the minimum idle time can be reduced to 3 minutes. Automatic Start Failure The EEC detects an automatic start failure if there are any of the following faults detected: •
Slow N2 spool-up
•
Rapidly increasing EGT
•
N1 roll back
•
N1 locked rotor: The EEC determines no N1 engine low-pressure rotor speed after sensing N2 has increased beyond 30%
•
EEC unable to command igniters: The EEC is unable to command either igniter in either channel
•
Loss of EGT indication
•
Inability to control fuel flow
When N2 is above 50%, the automatic start abort function is not available. The L(R) ENG switch must be moved to OFF to abort the engine start.
Auto Restart On ground, if the EEC senses an automatic start failure, the EEC turns the fuel and ignition OFF, and commands a dry motor to clear fuel from the engine for 30 seconds. After motoring the engine, the EEC automatically commands an auto restart. If the auto restart attempt fails, the EEC aborts the start. If the EEC detects an overspeed or a fuel metering valve fault, the EEC does not attempt an auto restart.
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TECHNICAL TRAINING MANUAL
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74-80-42
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
Automatic Start Begins Slow N2 Spool-Up
Yes
No
Yes
Overspeed Detected
Automatic Start Abort
Fail
Auto Restart
Successful
Engine Running
No Yes
Rapid EGT Rise or Loss of Signal
Yes
No
Fuel Metering Valve Operational
No
No
Engine Dry Spool 30 seconds
Yes No
N1 Roll Back
Yes
Igniters Operational
No
Yes
No EGT Signal
Automatic Start Continues
Yes
No N1 @ 30% N2
No
CS1_CS3_7400_015
Yes
Figure 20: Abnormal Starts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-43
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
Fuel Depulsing
Engine Start Limits
Fuel depulsing is an EEC function that uses the fuel metering valve (FMV) in the integrated fuel pump and control (IFPC) to turn fuel on and off, in an attempt to clear an impending surge or hot start.
The following table summarizes the engine starting limits.
The EEC performs fuel depulsing when a surge or an impending hot start, as indicated by a rapid rise in EGT or an exceedance of 1054°C (1929° F). If an impending hot start is detected, the EEC cycles fuel off for 2 seconds and back on for 12 seconds using the fuel metering valve. The EEC performs this cycling until EGT drops below the impending hot start detection limit or the surge has cleared. Fuel depulsing is inhibited whenever any of the following conditions are present: •
The engine N2 is above 50%
•
The aircraft is flying above an altitude of 20,000 ft
•
EGT signal is lost
Table 3: Engine Starting Limits START LIMITS Maximum EGT
1054°C (1929°F)
Minimum oil temperature for starting
-40°C (-40°F)
Minimum oil temperature for operation above idle
-6°C (21°F)
Maximum Oil Temperature Starter Duty Cycle
174°C (345°F) Three starter crank cycles for a total of 4 minutes cranking time maximum followed by 30 minutes of cooling time. 4 minutes maximum continuous cranking.
A timer in the EEC is used to discontinue fuel depulsing once the maximum allowable time of 28 seconds on the ground or 140 seconds in-flight has been reached, or if EGT increases to the starting amber value. The time limit is sufficient for two depulse cycles if the aircraft is on the ground. If the time limit has been exceeded, or the EGT has reached the starting amber value, the EEC terminates fuel depulsing and dry motors the engine for 30 seconds.
Start Abort
•
No N1 rotation by 30% N2
(EEC aborts start automatically below 54% N2)
•
N2 fails to reach stabilized idle speed within 120 seconds
•
No EGT rise after 20 seconds from positive fuel flow indication
•
Oil pressure does not increase within 30 seconds
•
EGT greater than 1054°C (Hot Start)
When the aircraft is in flight and below 20,000 ft, the timer in the EEC is changed to 140 seconds for fuel depulsing.
CS1_CS3_TB74-80_001
The EEC conducts a ground auto restart when the EGT has decreased below the impending hot start temperature, which the EEC determines depending on aircraft weight-on-wheels (WOW) status and N2 speed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-44
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
Impending Hot Start or Sub-Idle Compressor Stall N2 > 50%
Yes
Hot Start or Sub-Idle Compressor Stall Clear
Fuel Depulsing Inhibited
No
ALT > 20,000 ft
Continue Automatic Start
Yes
Dry Motor 30 seconds, then Auto Restart
Yes
Automatic start Abort and Dry Motor 30 seconds
No
Time Exceeded Automatic Start 28 seconds on Ground or 140 seconds in Flight
Fuel Depulsing Maximum 140 seconds
Yes
No
No
EGT signal is lost
Yes
Yes
Fuel Depulsing Maximum 28 seconds Yes
In Flight
No
Hot Start 1054 °C
CS1_CS3_7400_018
No
No
Figure 21: Fuel Depulsing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-45
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the starting and ignition system.
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TECHNICAL TRAINING MANUAL
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74-80-46
CS130_74-80_L2_Starting_MonitoringTests.fm
74-80 - Ignition and Starting 80-11 Starting
CAS MESSAGES
Table 6: INFO Messages Table 4: CAUTION Messages
MESSAGE
MESSAGE
80 L ENGINE INFO message will be set when one of the following is FAULT- STARTER true: SYSTEM INOP A dual channel starter speed signal fault, OR a dual channel SAV fault.
LOGIC
L ENG START ABORT
L engine starting procedure aborted.
R ENG START ABORT
R engine starting procedure aborted.
LOGIC
L ENG STARTER ATS is not disengaging or the L starter air valve (SAV) FAIL ON failed to open. R ENG STARTER ATS is not disengaging or the R starter air valve (SAV) FAIL ON failed to open.
A dual channel SAV failed closed fault has been detected by the right FADEC system. In the case of dual channel sensor fault, if the opposite channel control processor (CPU) has failed, a single channel fault would set this message.
Table 5: ADVISORY Messages MESSAGE
LOGIC
L (R) ENG FAULT Loss of redundant or non-critical function for the left or right engine. L (R) ENG STARTER OVHT
Left or right starter usage does not fulfill the duty cycle criteria: A 30 minute cool down time for 4 minutes (low N2), or 30 seconds (high N2).
ENGINE START DELAY
The engine is performing dry crank cooling during a ground start for protection against rotor bow.
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74-80-47
CS130_74-80_L2_Starting_MonitoringTests.fm
74-80 - Ignition and Starting 80-11 Starting
PRACTICAL ASPECTS STARTER OIL SERVICING The starter oil is changed at specific intervals as per the aircraft maintenance schedule. To service the oil, both the fill and overfill plugs are removed. The oil is added until it spills from the overfill plug, at which point the starter sight glass should read full. STARTER CHIP COLLECTOR The air turbine starter (ATS) chip collector can be removed by hand without draining the oil. A check valve in the fitting prevents the loss of oil when the starter chip collector is removed.
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TECHNICAL TRAINING MANUAL
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74-80-48
CS130_74-80_L2_Starting_PracticalAspects.fm
74-80 - Ignition and Starting 80-11 Starting
Fill Port
Sight Glass L FU
L
Overfill Port
Drain Port
Starter Chip Collector
CS1_CS3_7400_014
D AD
Figure 22: Starter Oil Servicing and Starter Chip Collector Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-49
CS130_74-80_L2_Starting_PracticalAspects.fm
74-80 - Ignition 80-11 Starting
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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74-80-50
CS130_74-80_ILB.fm
ATA 75 - Engine Air
BD-500-1A10 BD-500-1A11
75 - Engine Air
Table of Contents 75-24 Turbine Active Clearance Control System....................75-2 General Description .........................................................75-2 Component Location ........................................................75-4 Detailed Description .........................................................75-6 Monitoring and Tests .......................................................75-8 CAS Messages ............................................................75-9 75-25 Turbine Cooling Air ....................................................75-10 General Description .......................................................75-10 75-26 Buffer Air System........................................................75-12 General Description ........................................................75-12 Component Location ......................................................75-14 Buffer Air Heat Exchanger .........................................75-14 Buffer Air Shutoff Valve..............................................75-14 Buffer Air Shutoff Valve Solenoid...............................75-14 Buffer Air Pressure Sensor ........................................75-14 Buffer Air Check Valve...............................................75-14 Monitoring and Tests .....................................................75-16 CAS Messages ..........................................................75-17
High-Pressure Compressor Stator Vane Actuator .... 75-20 Detailed Description ...................................................... 75-22 Monitoring and Tests .................................................... 75-24 CAS Messages ......................................................... 75-25 75-32 Compressor Bleed Air System................................... 75-26 General Description ...................................................... 75-26 Low-Pressure Compressor Bleed Air System ........... 75-26 High-Pressure Compressor Bleed Air System .......... 75-26 Component Location ..................................................... 75-28 2.5 Bleed Air Valve ................................................... 75-28 2.5 Bleed Valve Actuator........................................... 75-28 High-Pressure Compressor Bleed Valve .................. 75-28 High-Pressure Compressor Bleed Valve Pressure Sensor ....................................................... 75-28 Detailed Description ..................................................... 75-30 Monitoring and Tests .................................................... 75-32 CAS Messages ......................................................... 75-33 Practical Aspects ........................................................... 75-34 Electronic Engine Control Actuator Test ................... 75-34
75-31 Variable Stator Vane System .....................................75-18 General Description .......................................................75-18 Low-Pressure Compressor ........................................75-18 High-Pressure Compressor .......................................75-18 Component Location ......................................................75-20 Low-Pressure Compressor Inlet Guide Vane Bellcrank and Synchronizing Ring .............................75-20 Low-Pressure Compressor Stator Vane Actuator......75-20 High-Pressure Stator Vane Bellcrank and Synchronizing Rings ...........................75-20 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-i
CS130_M5_75_EngineAirTOC.fm
75 - Engine Air
List of Figures Figure 1: Turbine Active Clearance Control System ............75-3 Figure 2: Turbine Active Clearance Control System Component Location .............................................75-5 Figure 3: Turbine Active Clearance Control System Schematic..............................................................75-7 Figure 4: Turbine Cooling Air..............................................75-11 Figure 5: Buffer Air System ................................................75-13 Figure 6: Buffer Air System Component Location ..............75-15 Figure 7: Variable Stator Vane System .............................75-19 Figure 8: Variable Stator Vane Control System Component Location ...........................................75-21 Figure 9: Variable Stator Vane Control Detailed Description ............................................75-23 Figure 10: Compressor Bleed Air System ............................75-27 Figure 11: Compressor Bleed Air System Component Location ...........................................75-29 Figure 12: Compressor Bleed Air System Schematic ..........75-31 Figure 13: Electronic Engine Control Actuator Test .............75-35
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75-ii
CS130_M5_75_EngineAirLOF.fm
75 - Engine Air
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TECHNICAL TRAINING MANUAL
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75-iii
CS130_M5_75_EngineAirLOF.fm
ENGINE AIR SYSTEMS - CHAPTER BREAKDOWN ENGINE AIR CHAPTER BREAKDOWN
Engine Cooling
1
Compressor Airflow
2
75 - Engine Air 75-24 Turbine Active Clearance Control System
75-24 TURBINE ACTIVE CLEARANCE CONTROL SYSTEM GENERAL DESCRIPTION The turbine active clearance control system provides fan air to the turbine cases to limit turbine case growth due to thermal expansion. This reduces high-pressure turbine (HPT) and low-pressure turbine (LPT) blade tip clearance and improves fuel efficiency. The turbine active clearance control (ACC) system consists of a fuel actuated turbine ACC air valve that regulates the fan airflow to the LPT and HPT cooling air manifolds. The cooling air manifolds spray air on the turbine case. The turbine ACC air valve is controlled by the electronic engine control (EEC). Servo fuel, supplied from the integrated fuel pump and control (IFPC) positions the valve in response to EEC commands.
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75-2
CS130_75_L2_TACCS_General.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
ELECTRONIC ENGINE CONTROL
Fan Air
TURBINE ACTIVE CLEARANCE CONTROL AIR VALVE
HIGH-PRESSURE TURBINE MANIFOLD
LEGEND
LOW-PRESSURE TURBINE MANIFOLD
INTEGRATED FUEL PUMP AND CONTROL
CS1_CS3_7524_005
Fan Air Servo Fuel Servo Fuel Return
Figure 1: Turbine Active Clearance Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-3
CS130_75_L2_TACCS_General.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
COMPONENT LOCATION The turbine active clearance control system has the following components: •
Turbine active clearance control air valve
•
High-pressure turbine (HPT) manifold
•
Low-pressure turbine (LPT) manifold
Turbine Active Clearance Control Air Valve The turbine active clearance control (ACC) air valve is located at the 1 o’clock position, near the rear of the high-pressure compressor (HPC). High-Pressure Turbine Manifold The high-pressure turbine (HPT) manifold is located on the HPT case. Low-Pressure Turbine Manifold The low-pressure turbine (LPT) manifold is located on the LPT case.
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75-4
CS130_75_L2_TACCS_ComponentLocation.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
LPT Manifold
Turbine Active Clearance Control Air Valve
CS1_CS3_7524_001
HPT Manifold
LEGEND HPT LPT
High-Pressure Turbine Low-Pressure Turbine
Figure 2: Turbine Active Clearance Control System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-5
CS130_75_L2_TACCS_ComponentLocation.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
DETAILED DESCRIPTION The turbine ACC air valve has a torque motor (T/M) that positions the valve using IFPC servo fuel pressure. The torque motor control is done by the EEC channel in control. The turbine ACC air valve single channel linear variable differential transformer (LVDT) sends position feedback to channel A of the EEC. The EEC controls the valve according to engine parameters and altitude in combination with the turbine ACC air valve position feedback. The turbine ACC air valve operation is scheduled as a function of N2 and altitude. The turbine ACC air valve is closed during takeoff and most of climb. It can open during the end of the climb phase and is always open in cruise. The failsafe mode of the turbine ACC air valve is the closed position.
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75-6
CS130_75_L3_TACCS_DetailedDescription.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
PMAG N2 SPEED
DMC 1 and DMC 2
Air Data
EEC Channel B
IFPC
Channel A
TURBINE ACC AIR VALVE T/M
LVDT
HPT MANIFOLD
Fan Bypass Air LEGEND Data Concentrator Unit Module Cabinet Electronic Engine Control High-Pressure Turbine Integrated Fuel Pump and Control Low-Pressure Turbine Linear Variable Differential Transformer Permanent Magnet Alternator Generator Torque Motor Fan Air Servo Fuel Servo Fuel Return
LPT MANIFOLD CS1_CS3_7524_004
DMC EEC HPT IFPC LPT LVDT PMAG TM
Figure 3: Turbine Active Clearance Control System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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75-7
CS130_75_L3_TACCS_DetailedDescription.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the active clearance control system.
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TECHNICAL TRAINING MANUAL
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75-8
CS130_75_L2_TACCS_MonitoringTests.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
CAS MESSAGES Table 1: ADVISORY Message MESSAGE L (R) ENG FAULT
LOGIC Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 2: INFO Messages MESSAGE
LOGIC
75 L ENGINE FAULT - ACC FAIL CLSD
INFO message is set when one of the following is true: • A dual-channel ACC torque motor fault exists • Single channel ACC feedback fault has been detected by the left FADEC system • The turbine ACC air valve stuck closed
75 R ENGINE FAULT - ACC FAIL CLSD
INFO message is set when one of the following is true: • A dual-channel ACC torque motor fault exists • Single channel ACC feedback fault has been detected by the right FADEC system • The turbine ACC air valve stuck closed
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75-9
CS130_75_L2_TACCS_MonitoringTests.fm
75 - Engine Air 75-25 Turbine Cooling Air
75-25 TURBINE COOLING AIR GENERAL DESCRIPTION The turbine cooling air system consists of a series of tubes that provide 4th or 6th stage high-pressure compressor (HPC) air to cool parts of the engine. The 6th stage HPC bleed air cools the following engine parts; •
High-pressure turbine (HPT) 2nd stage vanes
•
HPT 2nd stage blade attachment
The 4th stage HPC bleed air cools the following parts; •
Turbine intermediate case (TIC) fairing
•
Low-pressure turbine (LPT) case
•
LPT rotor and blade attachment
There are four tubes that feed cooling air from the 6th stage HPC to the HPT 2nd stage vanes and blade attachments. The air is metered by plates between the tubes and the HPT case. The cooling air then flows through hollow vanes and exits out the trailing edge. Seven TIC cooling air tubes cool the TIC fairings, inner walls and outer walls known as transition ducts. Four are supply tubes, and three are jumper tubes. The three jumper tubes feed air through the TIC fairings to the LPT rotors and blade attachment. Three additional jumper tubes feed air into the space between the LPT case and 2nd stage vane.
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75-10
CS130_75_L2_TCA_General.fm
75 - Engine Air 75-25 Turbine Cooling Air
HPT LPT
CS1_CS3_7525_001
LEGEND 6th Stage Air to HPT 4th Stage Air to LPT High-Pressure Turbine Low-Pressure Turbine
Figure 4: Turbine Cooling Air Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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75-11
CS130_75_L2_TCA_General.fm
75 - Engine Air 75-26 Buffer Air System
75-26 BUFFER AIR SYSTEM GENERAL DESCRIPTION The engine bearing compartments are cooled using a split buffer bearing cooling system. This supplies the bearing compartments with cooling and pressurizing air from the 2.5 low-pressure compressor (LPC) exit air, or from the 4th stage high-pressure compressor air (HPC), depending on engine operating conditions. The system is operated by the controlling electronic engine control (EEC) channel. The EEC controls the system through the buffer air valve solenoid (BAVS). The BAVS opens the buffer air shutoff valve (BASOV) when the BAVS is energized. The BASOV controls the flow of 4th stage air into the low-pressure side of the buffer air system to maintain the required seal pressure in each of the engine's bearing compartments during low power settings.
At low power, the 2.5 air does not have sufficient pressure and flow to supply the buffer air system. The EEC energizes the BAVS to open the BASOV. The BASOV allows 4th stage air to supply the front bearing compartments, consisting of the no. 1, no 1.5 and no. 2 bearings, no. 3, and no. 5/6 compartments, with air for cooling and pressurizing the seals. The no. 4 bearing compartment is always supplied with cooled 4th stage air regardless of the engine’s power settings. At high power, the 2.5 air is sufficient to supply the buffer air system. The high power switchover occurs at 75% N1. The EEC de-energizes the BAVS to close the BASOV. When the BASOV closes, the BACV opens, allowing 2.5 air to supply the front bearing compartment no.3 and compartments no.5/6 with air for cooling and pressurizing the seals.
A buffer air pressure sensor (BAPS) monitors the pressure downstream of the BASOV to confirm that the BASOV is in the correct position for the engine power setting. The BAPS provides a signal to each EEC channel. The buffer air check valve (BACV) allows 2.5 air to supply the buffer air system at high power settings when the BASOV is closed. The BACV closes when the BASOV is open and supplies 4th stage air to the buffer air system. When the BACV is closed it prevents backflow from the 4th stage into the engine. The buffer air heat exchanger (BAHX) cools stage 4 air used for the no. 4 bearing compartment cooling and pressurization. The BAHX uses 2.5 air to cool the 4th stage air.
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75-12
CS130_75_L2_BAS_General.fm
75 - Engine Air 75-26 Buffer Air System
N1 SPEED SENSOR
BAHX
EEC BAVS
BASOV
SOL
BACV
No. 3 Bearing
BAPS
No. 5 Bearing
No. 6 Bearing
No. 5 and No. 6 Bearing Compartment
Forward Bearing Compartment
No. 3 Bearing Compartment
No. 4 Bearing Compartment
NOTE No. 1 Bearing
No. 1.5 Bearing
No. 2 Bearing
Engine shown at high power.
BACV BAHX BAPS BASOV BAVS EEC
2.5 Bleed Air 4th Stage Air Buffer Air Check Valve Buffer Air Heat Exchanger Buffer Air Pressure Sensor Buffer Air Shutoff Valve Buffer Air Valve Solenoid Electronic Engine Control
CS1_CS3_7526_001
LEGEND
Figure 5: Buffer Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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75-13
CS130_75_L2_BAS_General.fm
75 - Engine Air 75-26 Buffer Air System
COMPONENT LOCATION The buffer air cooling system consists of the following components: •
Buffer air heat exchanger (BAHX)
•
Buffer air shutoff valve (BASOV)
•
Buffer air shutoff valve solenoid (BAVS)
•
Buffer air pressure sensor (BAPS)
•
Buffer air check valve (BACV)
BUFFER AIR HEAT EXCHANGER The buffer air heat exchanger (BAHX) is located on the fan intermediate case in the 2.5 bleed cavity at the 12:00 o’clock position. BUFFER AIR SHUTOFF VALVE The buffer air shutoff valve (BASOV) is located on the rear of the high-pressure compressor (HPC) at the 11 o’clock position. BUFFER AIR SHUTOFF VALVE SOLENOID The buffer air shutoff valve solenoid (BAVS) is located on the right side of the HPC case at the 1 o’clock position. BUFFER AIR PRESSURE SENSOR The buffer air pressure sensor (BAPS) is located below the BAHX at the 2 o’clock position. BUFFER AIR CHECK VALVE The buffer air check valve (BACV) is located at the 12 o’clock position on the fan intermediate case.
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TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-14
CS130_75_L2_BAS_ComponentLocation.fm
75 - Engine Air 75-26 Buffer Air System
Buffer Air Heat Exchanger Buffer Air Check Valve
Gasket
A B C
A
C BUFFER AIR SHUTOFF VALVE SOLENOID
B
BUFFER AIR PRESSURE SENSOR
BUFFER AIR D SHUTOFF VALVE
CS1_CS3_7526_002
D
Figure 6: Buffer Air System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-15
CS130_75_L2_BAS_ComponentLocation.fm
75 - Engine Air 75-26 Buffer Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the buffer air system operations.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-16
CS130_75_L2_BAS_MonitoringTests.fm
75 - Engine Air 75-26 Buffer Air System
CAS MESSAGES Table 3: ADVISORY Message MESSAGE
LOGIC
L (R) ENG FAULT Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 4: INFO Messages MESSAGE
LOGIC
75 L (R) ENGINE FAULT - BAV INOP
INFO message is set by the left (right) engine when the buffer air shutoff valve is failed closed.
75 L (R) ENGINE FAULT - BACV FAIL CLSD
INFO message is set by the left (right) engine when the buffer air check valve is failed closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-17
CS130_75_L2_BAS_MonitoringTests.fm
75 - Engine Air 75-31 Variable Stator Vane System
75-31 VARIABLE STATOR VANE SYSTEM GENERAL DESCRIPTION The variable stator vane system adjusts the position of the low-pressure compressor (LPC) inlet guide vanes and the high-pressure compressor (HPC) variable stator vanes during engine operation. This optimizes engine performance. LOW-PRESSURE COMPRESSOR
The HPC stator vane actuator moves the HPC inlet guide vane and the stator vanes of the first three stages mechanically through a bellcrank and synchronizing ring. The HPC stator vanes move between the open and closed position based on N2 speed. The HPC variable stator vanes are closed during engine startup and engine idle, and are fully open at takeoff power. The actuators require no rigging during installation. They are designed to fail in the open position to allow maximum airflow through the compressors.
The electronic engine control (EEC) controls the LPC stator vane actuator position through a torque motor. This controls a servovalve mounted on the actuator. A linear variable differential transformer (LVDT) provides feedback to the EEC. The LPC stator vane actuator is fuel actuated using servo fuel supplied by the integrated fuel pump and control (IFPC). The LPC stator vane actuator moves the LPC inlet guide vane mechanically through a bellcrank and synchronizing ring. The LPC inlet guide vane moves between the open and closed position, based on N1 speed. The LPC inlet guide vanes move toward the closed position during engine start up and idle, modulates during transient operations, and are open above idle power. HIGH-PRESSURE COMPRESSOR The EEC controls the HPC stator vane actuator position through a torque motor. This controls a servovalve mounted on the actuator. A LVDT provides feedback to the EEC. The HPC stator vane actuator is fuel actuated using servo fuel, supplied by the IFPC.
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75-18
CS130_75_L2_VSVS_General.fm
75 - Engine Air 75-31 Variable Stator Vane System
N1 SPEED SENSOR
LPC SVA
EEC
IFPC
PMAG N2 SPEED
HPC SVA
BELLCRANK and SYNCHRONIZING RING
LPC INLET GUIDE VANES
BELLCRANK and SYNCHRONIZING RINGS
HPC INLET GUIDE VANE and STATOR VANES
LEGEND
CS1_CS3_7531_006
EEC HPC IFPC LPC PMAG SVA
Servo Fuel Servo Fuel Return Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Control Low-Pressure Compressor Permanent Magnet Alternator Generator Stator Vane Actuator Mechanical Link
Figure 7: Variable Stator Vane System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-19
CS130_75_L2_VSVS_General.fm
75 - Engine Air 75-31 Variable Stator Vane System
COMPONENT LOCATION The variable stator vane control system consists of the following components: •
LPC inlet guide vane bellcrank and synchronizing ring
•
LPC stator vane actuator
•
HPC stator vane bellcrank and synchronizing rings
•
HPC stator vane actuator
LOW-PRESSURE COMPRESSOR INLET GUIDE VANE BELLCRANK AND SYNCHRONIZING RING The LPC inlet guide vane (IGV) bellcrank and synchronizing ring is mounted around the engine at the inlet of the LPC. LOW-PRESSURE COMPRESSOR STATOR VANE ACTUATOR The LPC SVA is mounted at the 5 o’clock position on the compressor intermediate case (CIC) firewall. HIGH-PRESSURE STATOR VANE BELLCRANK AND SYNCHRONIZING RINGS The HPC stator vane bellcrank and synchronizing rings are mounted on the HPC. HIGH-PRESSURE COMPRESSOR STATOR VANE ACTUATOR The HPC stator vane actuator is mounted on the HPC case on the right side of the engine at the 4 o’clock position.
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TECHNICAL TRAINING MANUAL
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75-20
CS130_75_L2_VSVS_ComponentLocation.fm
75 - Engine Air 75-31 Variable Stator Vane System
Low Pressure Compressor Synchronizing Ring
High Pressure Compressor Bellcrank High Pressure Compressor Starter Vane Actuator Low Pressure Compressor Starter Vane Actuator
Low Pressure Compressor Bellcrank
CS1_CS3_7531_001
High Pressure Compressor Synchronizing Rings
Figure 8: Variable Stator Vane Control System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-21
CS130_75_L2_VSVS_ComponentLocation.fm
75 - Engine Air 75-31 Variable Stator Vane System
DETAILED DESCRIPTION Either channel of the EEC controls the LPC SVA through a dual-coil torque motor which controls an internal servovalve. The position of the LPC SVA is sensed by a dual-coil linear variable differential transformer (LVDT). The LVDT is used to transmit the piston position to each electronic engine control (EEC) channel. If there is a power failure to the LPC SVA, the actuator positions the inlet guide vanes in the open position for maximum airflow through the LPC. The LPC SVA is positioned based on the N1 speed sensor signal. Either channel of the EEC control the HPC SVA through a dual-coil torque motor that controls an internal servovalve. The position of the HPC SVA is sensed by a dual-coil LVDT. The LVDT is used to transmit the piston position to each EEC channel. If there is a power failure to the LPC SVA, the actuator positions the inlet guide vanes and stator vanes in the open position for maximum airflow through the HPC. The HPC SVA is positioned based on the N2 speed sensing information received from the permanent magnet alternator generator (PMAG).
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75-22
CS130_75_L3_VSVS_DetailedDescription.fm
75 - Engine Air 75-31 Variable Stator Vane System
N1 SPEED SENSOR
LPC STATOR VANE ACTUATOR
LPC
TORQUE MOTOR
BELLCRANK and SYNCHRONIZING RING
Channel A Channel B
Inlet Guide Vanes
LVDT LEGEND
Channel A EEC HPC IFPC LPC LVDT PMAG
Channel B EEC Channel A
IFPC
Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Control Low-Pressure Compressor Linear Variable Differential Transformer Permanent Magnet Alternator Generator Mechanical Link Servo Fuel Servo Fuel Return
Channel B HPC HPC STATOR VANE ACTUATOR
Inlet Guide Vanes
TORQUE MOTOR Channel A
LVDT
BELLCRANK and SYNCHRONIZING RINGS
Channel A PMAG N2 SPEED
1st Stage Stator Vanes
2nd Stage Stator Vanes
Channel B 3rd Stage Stator Vanes
CS1_CS3_7531_005
Channel B
Figure 9: Variable Stator Vane Control Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-23
CS130_75_L3_VSVS_DetailedDescription.fm
75 - Engine Air 75-31 Variable Stator Vane System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for air system operations.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-24
CS130_75_L2_VSVS_MonitoringTests.fm
75 - Engine Air 75-31 Variable Stator Vane System
CAS MESSAGES Table 5: CAUTION Messages MESSAGE
LOGIC
L ENG OPER DEGRADED
Uncertainty on T2/TAT (degraded or default value), or HPC stator vane actuator not tracking as it should.
R ENG OPER DEGRADED
Uncertainty on T2/TAT (degraded or default value), or HPC stator vane actuator not tracking as it should.
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TECHNICAL TRAINING MANUAL
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75-25
CS130_75_L2_VSVS_MonitoringTests.fm
75 - Engine Air 75-32 Compressor Bleed Air System
75-32 COMPRESSOR BLEED AIR SYSTEM GENERAL DESCRIPTION The engine has a compressor bleed air system (BAS) to provide greater compressor stability during starting and engine transient operation. Air is bled from the rear of the low-pressure compressor (LPC) at station 2.5, and from one 6th stage bleed port in the high-pressure compressor (HPC).
In the event of an engine surge, the EEC:
LOW-PRESSURE COMPRESSOR BLEED AIR SYSTEM
When the surge detection signal clears, bleed valve and the variable stator vanes are returned to their normal positions, and continuous ignition is turned off after a timer expires.
The low-pressure compressor BAS uses a bleed valve. It is located at the exit of the LPC at station 2.5, and provides an improved surge margin during starting, low power, and transient operation. The bleed valve forms a ring around the LPC case. The bleed valve extracts dirt, rain, and ice during ground operation. Air from the bleed valve is vented to the fan air stream. The 2.5 bleed actuator has an integral torque motor assembly that controls servo fuel pressure from the integrated fuel pump and control (IFPC) to position the actuator. The 2.5 bleed valve is attached to the LPC just forward of the compressor intermediate case (CIC). The actuator acts on the bellcrank to open and close the valve. The valve is fully modulated through the electronic engine control (EEC) commands. The 2.5 bleed valve actuator is scheduled based on N1 speed and air data information. The 2.5 bleed valve actuator moves the 2.5 bleed valve ring to the open position during startup and idle, modulates during transient operations, and closes at takeoff. The failsafe position of this bleed is in the closed position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
•
Opens the bleed valve
•
Activates continuous ignition on both igniters
•
Resets the variable stator vanes to recover from the surge condition.
HIGH-PRESSURE COMPRESSOR BLEED AIR SYSTEM The high-pressure compressor (HPC) BAS discharges the 6th stage compressor air into the engine core area. The HPC bleed valve is augmented by the cowl anti-ice valves (CAIVs), which open during engine start and discharges air through the cowl anti-ice (CAI) ducting. For more information on the cowl anti-ice system (CAIS), refer to ATA 30 Ice and Rain Protection. The HPC bleed valve is a two-position poppet-type valve. The valve is spring-loaded to the open position during engine starting. The valve is closed using station 2.9 HPC air. The valve closes when HPC pressure is high enough to overcome the spring-force, and closes fully at idle power. The HPC bleed valve sensor provides the EEC with feedback on the position of the HPC bleed valve.
TECHNICAL TRAINING MANUAL
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75-26
CS130_75_L2_CBAS_General.fm
75 - Engine Air 75-32 Compressor Bleed Air System
HPC BLEED VALVE
HPC BLEED VALVE PRESSURE SENSOR
CAI VALVES
IFPC
2.5 BLEED VALVE
2.5 BLEED ACTUATOR
EEC
PMAG N2 SPEED
Air Data
CAI EEC HPC IFPC PMAG
Bleed Air Servo Fuel Servo Fuel Return Cowl Anti-Ice Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Control Permanent Magnet Alternator Generator Mechanical Link
NOTE 2.5 bleed valve position is based on N1 speed.
CS1_CS3_7532_006
LEGEND
Figure 10: Compressor Bleed Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-27
CS130_75_L2_CBAS_General.fm
75 - Engine Air 75-32 Compressor Bleed Air System
COMPONENT LOCATION The compressor bleed air system consists of the following components: •
2.5 bleed air valve
•
2.5 bleed valve actuator
•
High-pressure compressor (HPC) bleed valve
•
High-pressure compressor (HPC) bleed valve pressure sensor
2.5 BLEED AIR VALVE The 2.5 bleed air valve is located between the low-pressure compressor (LPC) and the high-pressure compressor (HPC). 2.5 BLEED VALVE ACTUATOR The 2.5 bleed valve actuator is located on the rear of the compressor intermediate case fireseal at the 3 o’clock position. HIGH-PRESSURE COMPRESSOR BLEED VALVE The high-pressure compressor (HPC) bleed valve is located on the HPC case at the 1 o’clock position. HIGH-PRESSURE COMPRESSOR BLEED VALVE PRESSURE SENSOR The high-pressure compressor (HPC) bleed valve pressure sensor is located at the 2 o’clock position on the compressor intermediate case fireseal.
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75-28
CS130_75_L2_CBAS_ComponentLocation.fm
75 - Engine Air 75-32 Compressor Bleed Air System
2.5 Bleed Valve Actuator
2.5 Bleed Valve
CS1_CS3_7532_001
High Pressure Compressor Bleed Valve
High Pressure Compressor Bleed Valve Pressure Sensor
Figure 11: Compressor Bleed Air System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-29
CS130_75_L2_CBAS_ComponentLocation.fm
75 - Engine Air 75-32 Compressor Bleed Air System
DETAILED DESCRIPTION A dual-channel LVDT, mounted in the valve actuator, provides feedback to the EEC proportional to the position of the piston. The EEC directs the fuel pressure supplied through a dual-coil torque servomotor system. This drives the valve to the required position. Scheduling of the 2.5 bleed valve actuator is based on N1 speed. For steady-state operation, the bleed valve is modulated between the fully opened and fully closed positions. The bleed valve position is scheduled as a function of N1 biased with altitude and Mach number. At takeoff power conditions, the bleed valve is fully closed. For starting and accelerations, the bleed valve follows the steady-state schedule. For decelerations, the bleed valve opens to protect LPC stability. In the event of an engine surge, the bleed valve opens to enhance recovery.
The HPC bleed valve pressure sensor sends a signal to the EEC to determine if the HPC bleed valve has failed in the open position, exposing the engine core to hot compressor air after the engine start. The sensor is also used by the EEC to determine if the HPC bleed valve has failed to close during engine start. If the valve is stuck closed during the engine start, the HPC could be overloaded, which could lead to an engine stall.
Upon detection of a surge during takeoff, the EEC assumes foreign object debris (FOD) and the bleed valve is partially opened to allow a 15% bleed flow. Assuming there is no other damage, this results in normal engine parameters except for a higher exhaust gas temperature (EGT) in the range of 45°C. The bleed valve remains in this position until a power setting below approach idle is selected, or the aircraft is not in the takeoff configuration. When the surge clears, the 2.5 bleed is positioned based on the normal control scheduling, and EGT is restored to normal levels. During engine start, the HPC bleed valve is open to unload the HPC. As the engine speed increases, the HPC moves to the closed position. In addition to the HPC bleed valve, the EEC also commands the cowl anti-ice valves (CAIV) open during start.
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TECHNICAL TRAINING MANUAL
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75-30
CS130_75_L3_CBAS_DetailedDescription.fm
75 - Engine Air 75-32 Compressor Bleed Air System
IFPC CAI VALVES
Torque Motor N1 SPEED SENSOR
Channel A
EEC
2.5 BLEED VALVE
Channel B Channel A 2.5 BLEED ACTUATOR
Air Data • Altitude • Mach Number
BELLCRANK
Channel B LVDT Channel A
Channel B LEGEND HPC PRESSURE SENSOR
Channel A HPC BLEED VALVE Channel B
NOTE Left engine shown. Right engine similar.
CS1_CS3_7532_005
CAI EEC HPC IFPC LVDT
HPC Bleed Air Servo Fuel Servo Fuel Return Cowl Anti-Ice Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Controller Linear Variable Differential Transformer Mechanical Link
Figure 12: Compressor Bleed Air System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-31
CS130_75_L3_CBAS_DetailedDescription.fm
75 - Engine Air 75-32 Compressor Bleed Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the air system operations.
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TECHNICAL TRAINING MANUAL
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75-32
CS130_75_L2_CBAS_MonitoringTests.fm
75 - Engine Air 75-32 Compressor Bleed Air System
CAS MESSAGES Table 6: CAUTION Messages MESSAGE
LOGIC
L ENG NACELLE OVHT
Left burst duct or L HPC valve failed open, and automation failed to close it.
R ENG NACELLE OVHT
Right burst duct or R HPC valve failed open, and automation failed to close it.
Table 7: ADVISORY Message MESSAGE L (R) ENG FAULT
LOGIC Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 8: INFO Message MESSAGE 75 L (R) ENGINE FAULT - HPC BLEED VLV INP
LOGIC INFO message is set by the left (right) engine when the EEC active channel has detected that the HPC bleed valve has failed to close.
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75-33
CS130_75_L2_CBAS_MonitoringTests.fm
75 - Engine Air 75-32 Compressor Bleed Air System
PRACTICAL ASPECTS ELECTRONIC ENGINE CONTROL ACTUATOR TEST The actuator test detects faults on any of the fuel controlled actuators. The test requires the engine to be dry motored. The electronic engine control (EEC) performs interactive tests on the following components: •
The fuel metering valve in the IFPC
•
The low-pressure compressor (LPC) variable stator vane actuator
•
The high-pressure compressor (HPC) variable stator vane actuator
•
The low-pressure compressor (LPC) 2.5 bleed valve actuator
•
The turbine active clearance control (ACC) actuator
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75-34
CS130_75_L2_CBAS_PracticalAspects.fm
75 - Engine Air 75-32 Compressor Bleed Air System
MAINT MENU EEC1-A-ACTUATOR TEST
1/5
*** WARNING *** ENGINE WILL DRY CRANK / DRY MOTOR AND COMPONENTS WILL MOVE DURING TEST EDP WILL BE PRESSURIZED AND PMAG WILL BE ENERGIZED
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- FADEC POWER IS APPLIED 3- FUEL IS PRESENT IN ENGINE 4- AIR PRESSURE IS AVAILABLE AT STARTER INLET
* NOTE *
PRESS "CONTINUE" TO GO TO START PAGE
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7321_019
TEST DURATION: APPROXIMATELY 150 SECONDS
Figure 13: Electronic Engine Control Actuator Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-35
CS130_75_L2_CBAS_PracticalAspects.fm
75 - Engine Air 75-32 Compressor Bleed Air System
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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75-36
CS130_75_ILB.fm
ATA 76 - Engine Controls
BD-500-1A10 BD-500-1A11
76 - Engine Controls
Table of Contents 76-10 Engine Control..............................................................76-2 General Description ..........................................................76-2 Component Location ........................................................76-4 Component Information ....................................................76-6 Throttle Quadrant Assembly ........................................76-6 Detailed Description .........................................................76-8 Engine Control .............................................................76-8 Thrust Reverser Control...............................................76-8 Autothrottle Control ......................................................76-8 Thrust Lever Warning ..................................................76-8 Monitoring and Tests ......................................................76-10 CAS Messages ..........................................................76-11
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76-i
CS130_M5_76_EngineCtlsTOC.fm
76 - Engine Controls
List of Figures Figure 1: Engine Control System..........................................76-3 Figure 2: Throttle Quadrant Assembly Component Location .............................................76-5 Figure 3: Throttle Quadrant Assembly..................................76-7 Figure 4: Engine Control System Schematic........................76-9
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76-ii
CS130_M5_76_EngineCtlsLOF.fm
76 - Engine Controls
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76-iii
CS130_M5_76_EngineCtlsLOF.fm
ENGINE CONTROLS - CHAPTER BREAKDOWN ENGINE CONTROLS CHAPTER BREAKDOWN
Throttle Quadrant Assembly
1
76 - Engine Controls 76-10 Engine Control
76-10 ENGINE CONTROL GENERAL DESCRIPTION Engine control is carried out using thrust levers that are located on the throttle quadrant assembly (TQA). Forward and reverse thrust control for each engine is provided by a single lever. Engine RUN switches, also located on the TQA, are used for commanding engine starts. The left and right lever systems of the TQA are completely independent of each other. Each thrust lever supplies thrust command signals to the electronic engine control (EEC). Thrust lever position information is sent to the autothrottle system in the data concentrator unit module cabinets (DMCs), through the brake data concentrator unit (BDCU). Thrust command signals are used by the EEC to calculate the fuel scheduling required to achieve the desired thrust. The EEC then outputs a command to the integrated fuel pump and control (IFPC) to schedule the fuel flow. When starting the engine, the engine RUN switches on the TQA are moved to the RUN position. A signal is sent to the EEC to command the starter air valve (SAV), IFPC, and ignition system in the start sequence. Engine shutdown is initiated by moving the same engine RUN switch to OFF. The thrust levers are either controlled automatically by the autothrottle system, or manually. Thrust reversers (T/Rs) are deployed by moving the thrust levers into the reverse thrust range. Thrust reverser stow/deploy switches in the TQA detect that thrust reverse has been selected and supply signals to unlock the track lock units.
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76-2
CS130_76_L2_EngineControl_General.fm
76 - Engine Controls 76-10 Engine Control
THROTTLE QUADRANT ASSEMBLY IGNITION EXCITER BOX DMC MAX
MAX
IDLE
IDLE
Autothrottle
ELECTRONIC ENGINE CONTROL
INTEGRATED FUEL PUMP AND CONTROL
BRAKE DATA CONCENTRATOR UNIT L ENG
R ENG
OFF
OFF
ON
ON
STARTER AIR VALVE
THRUST REVERSER TRACK LOCK UNIT
DMC
CS1_CS3_7610_001
LEGEND Data Concentrator Unit Module Cabinet ARINC 429
Figure 1: Engine Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-3
CS130_76_L2_EngineControl_General.fm
76 - Engine Controls 76-10 Engine Control
COMPONENT LOCATION The engine throttle quadrant assembly (TQA) is located on the fight deck center pedestal.
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76-4
CS130_76_L2_EngineControl_ComponentLocation.fm
76 - Engine Controls
MAX
MAX
IDLE
IDLE
REV
REV
MAX REV
MAX REV
L ENG
R ENG
ON
ON
OFF
OFF
CS1_CS3_7600_015
76-10 Engine Control
Figure 2: Throttle Quadrant Assembly Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-5
CS130_76_L2_EngineControl_ComponentLocation.fm
76 - Engine Controls 76-10 Engine Control
COMPONENT INFORMATION
Rotary Variable Differential Transformers
THROTTLE QUADRANT ASSEMBLY
There are two types of rotary variable differential transformers (RVDTs) in the TQA, dual and single channel. Each thrust lever is connected to one of each through mechanical gears inside the TQA.
The throttle quadrant assembly (TQA) is a line replaceable unit (LRU) that houses both engine thrust levers and the L(R) ENG switches. The components that are part of the TQA are:
Baulk Solenoids The baulk solenoids prevent the thrust levers from moving beyond the reverse idle position.
•
Thrust levers
•
ENG switches
•
RVDTs
Thrust Reverser Stow/Deploy Switches
•
Baulk solenoids
•
Thrust reverser stow/deploy switches
•
Autothrottle controller
The thrust reverser stow/deploy switches are single pole double throw type. They are activated when the thrust levers move into the reverse range.
•
Autothrottle servomotors
Autothrottle Controller The autothrottle controller provides output to the autothrottle servomotor.
Thrust Levers The thrust levers travel a total of 67° from the MAX REV to MAX positions. the thrust levers have hard stops at three positions; IDLE, MAX and MAX REV. A friction clutch provides a feel force and maintains the thrust lever in the selected position. The levers have an autothrottle disconnect takeoff/go-around (TOGA) switch built in.
switch
and
Autothrottle Servomotors Each thrust lever has an autothrottle servomotor that moves the thrust levers in response to the autothrottle controller commands.
a
Each thrust lever has a finger lift that allows the lever to move from idle to the reverse thrust position. ENG Switches The L(R) ENG switches initiate the engine start during an automatic start. During a manual start, they command fuel and ignition on. The L(R) ENG switches are selected OFF to shut down the engines.
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76-6
CS130_76_L2_EngineControl_ComponentInformation.fm
76 - Engine Controls 76-10 Engine Control
67° Autothrottle Disconnect Switches Finger Lifts
Takeoff/Go-Around Switches Baulk Solenoid
Thrust Reverser Stow/Deploy Switches
MAX IDLE
Servomotor REV
ENG Switches MAX REV Autothrottle Controller CS1_CS3_7600_003
Rotary Variable Displacement Transformers
Figure 3: Throttle Quadrant Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-7
CS130_76_L2_EngineControl_ComponentInformation.fm
76 - Engine Controls 76-10 Engine Control
DETAILED DESCRIPTION ENGINE CONTROL Each throttle lever mechanically drives three electrically independent RVDTs housed into one dual-coil and one single coil housing. Excitation voltage for the dual RVDTs is provided by the electronic engine control (EEC). Excitation voltage for the single RVDT is provided by the brake data concentrator unit (BDCU). The dual-coil RVDTs provide thrust lever position signal to the EEC. Channel A of the RVDT provides a signal for channel A of the EEC and channel B of the RVDT provides a signal for channel B of the EEC. The third RVDT (single coil housing) interfaces with the BDCU to provide electrically independent, digitized thrust lever angle information. This information is available to all systems, including the power plant, through the data concentrator unit module cabinet (DMC). A L (R) THROTTLE FAIL caution message is displayed when the EEC cannot determine the throttle position. The engine thrust is reduced to idle and the engine does not respond to thrust lever movement. THRUST REVERSER CONTROL The EEC provides a signal to control the thrust reverser (T/R) system baulk solenoid. When the T/Rs are deployed 85% of their travel, the EEC releases the baulk solenoid. This releases the thrust lever and allows full reverse thrust to be applied. The TQA baulk solenoids receive power from the aircraft control and distribution cabinets (CDCs). The EEC provides both the logic and the ground for the solenoids operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The interlock, provided by the baulk solenoids, incorporates an override operation in both directions that requires 25 lb of force per lever to overcome. The TQA has two stow/deploy microswitches per lever that control 28 VDC power to the thrust reverser track lock solenoids by providing the ground in circuit. These switches are hardwired to their respective control and distribution cabinets (CDCs). If either the left or right thrust lever is moved into the reverse thrust range while the aircraft is in the air, a THROTTLE IN REVERSE caution message is displayed. AUTOTHROTTLE CONTROL The autothrottle controller receives power from the flight control panel (FCP) when autothrottle is engaged. When engaged, the autothrottle software sends throttle control information to the autothrottle controller inside the TQA. The autothrottle controller then sends the commands to the autothrottle servomotors inside the TQA to move the thrust levers. The thrust levers are equipped with an autothrottle disconnect switch to disengage the autothrottle. THRUST LEVER WARNING If the thrust lever is not at idle when the engine is being started, the EEC initiates an aural warning to warn the crew, but does not prevent the engine from starting. The aural warning sounds “THRUST LEVER” three times.
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76-8
CS130_76_L3_EngineControl_DetailedDescription.fm
76 - Engine Controls 76-10 Engine Control
TQA
DMC 1 and DMC 2
LEGEND BDCU CDC DMC EEC FCP IFPC RIU RVDT TQA T/R
Autothrottle Disconnect Switch
Autothrottle
FCP Autothrottle Controller
28 VDC
Autothrottle Servomotor Position Signal RIU 1 and RIU 2
BDCU 1
Excitation Voltage
Single Channel RVDT DUAL-CHANNEL RVDT Position Signal Channel A
Brake Data Concentrator Unit Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Engine Control Flight Control Panel Integrated Fuel Pump and Control Radio Interface Unit Rotary Variable Differential Transformer Throttle Quadrant Assembly Thrust Reverser ARINC 429 Discrete
EEC Channel A
Excitation Voltage
IFPC
Position Signal Channel B CDC 3
Channel B
Excitation Voltage
BAULK SOLENOID
TQA L REV LOCK DC SSPC 3A BUS 1
Ground
Logic: Always On “THRUST LEVER, THRUST LEVER, THRUST LEVER”
L Engine Run Switch T/R Stow/ Deploy Switches
28 VDC
NOTE Left engine throttle operation shown. Right engine throttle operation similar.
To Track Lock Unit Solenoids
CS1_CS3_7600_011
CDC 1
Figure 4: Engine Control System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-9
CS130_76_L3_EngineControl_DetailedDescription.fm
76 - Engine Controls 76-10 Engine Control
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the engine control system.
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TECHNICAL TRAINING MANUAL
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76-10
CS130_76_L2_EngineControl_MonitoringTests.fm
76 - Engine Controls 76-10 Engine Control
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L THROTTLE FAIL
Left throttle position is not recognized by the EEC.
R THROTTLE FAIL
Right throttle position is not recognized by the EEC.
THROTTLE IN REVERSE
Left or right thrust reverser selected in flight.
Table 2: ADVISORY Messages MESSAGE
LOGIC
L ENGINE FAULT Loss of redundant or non-critical function for the left engine. R ENGINE FAULT Loss of redundant or non-critical function for the right engine.
Table 3: STATUS Messages MESSAGE
LOGIC
L ENG SHUTDOWN
In-flight pilot commanded engine shut-down. Left engine run switch to OFF).
R ENG SHUTDOWN
In-flight pilot commanded engine shut-down. Right engine run switch to OFF.
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76-11
CS130_76_L2_EngineControl_MonitoringTests.fm
76 - Engine Controls 76-10 Engine Control
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TECHNICAL TRAINING MANUAL
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76-12
CS130_76_ILB.fm
ATA 77 - Engine Indicating
BD-500-1A10 BD-500-1A11
77 - Engine Indicating
Table of Contents 77-11 Engine Power Indicating System..................................77-2 General Description ..........................................................77-2 Component Location ........................................................77-4 Nf Speed Sensor..........................................................77-4 N1 Speed Probe ..........................................................77-4 Permanent Magnet Alternator Generator.....................77-4 Component Information ....................................................77-6 N1 Speed Probe ..........................................................77-6 Permanent Magnet Alternator Generator.....................77-8 Controls and Indications .................................................77-10 N1 Indication ..............................................................77-10 N1 Thrust Reference Bug ..........................................77-12 N1 Thrust Reference..................................................77-12 N1 Thrust Command Cursor......................................77-12 Takeoff Reference without FLEX or MAX Climb........77-12 N2 Indication ..............................................................77-14 Detailed Description .......................................................77-16 Monitoring and Tests .....................................................77-18 CAS Messages ..........................................................77-19
Monitoring and Tests .................................................... 77-28 CAS Messages ......................................................... 77-29 77-32 Vibration Monitoring................................................... 77-30 General Description ....................................................... 77-30 Component Location ..................................................... 77-32 Forward Vibration Sensor ......................................... 77-32 Aft Vibration Sensor .................................................. 77-32 Prognostics Health and Management Unit ............... 77-32 Data Storage Unit ..................................................... 77-32 Controls and Indications ................................................ 77-34 Indications ................................................................. 77-34 Detailed Description ...................................................... 77-36 Monitoring and Tests .................................................... 77-38 CAS Messages ......................................................... 77-39 Practical Aspects ............................................................ 77-40 Onboard Maintenance System.................................. 77-40 Fan Trim Balance Using the PW 1500G Trim Balance Helper ................................................. 77-58
77-21 Temperature Indicating System..................................77-20 General Description ........................................................77-20 Component Location ......................................................77-22 Exhaust Gas Temperature Probe and Cable Assembly .........................................................77-22 Oil Temperature Probe ..............................................77-22 Probe Junction ...........................................................77-22 Controls and Indications .................................................77-24 Indications..................................................................77-24 Detailed Description .......................................................77-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-i
CS130_M5_77_EngineIndicatingTOC.fm
77 - Engine Indicating
List of Figures Figure 1: Engine Power Indicating System...........................77-3 Figure 2: Engine Power Indicating System Component Location .............................................77-5 Figure 3: N1 Speed Probe....................................................77-7 Figure 4: Permanent Magnet Alternator Generator .............77-9 Figure 5: N1 Indication .......................................................77-11 Figure 6: N1 Thrust Reference Bug, N1 Thrust Reference, N1 Thrust Command Cursor, and Takeoff Reference........................................77-13 Figure 7: N2 Indication .......................................................77-15 Figure 8: Engine Power Indicating System Detailed Description ............................................77-17 Figure 9: Exhaust Gas Temperature Indicating System ................................................77-21 Figure 10: Exhaust Gas Temperature Indicating System Component Location..............77-23 Figure 11: Exhaust Gas Temperature Indications ................77-25 Figure 12: Exhaust Gas Temperature Detailed Description ............................................77-27 Figure 13: Vibration Monitoring ............................................77-31 Figure 14: Vibration Monitoring Component Location .........77-33 Figure 15: Vibration Monitoring Indications ..........................77-35 Figure 16: Vibration Monitoring Detailed Description ...........77-37 Figure 17: Trim Balance .......................................................77-41 Figure 18: Fan Hub Balance Weights...................................77-43 Figure 19: Trim Balance Procedure Main Menu...................77-45 Figure 20: Review VIB Data History .....................................77-47 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 21: Figure 22: Figure 23: Figure 24: Figure 25: Figure 26: Figure 27: Figure 28: Figure 29: Figure 30: Figure 31:
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Review/Edit Stored Weights ............................... 77-49 Perform Balance ................................................. 77-51 Display Balance Coefficients .............................. 77-53 Engine Serial Number Coefficients..................... 77-55 Modify Settings Screen....................................... 77-57 Fan Trim Balance ............................................... 77-59 Onboard Maintenance System Data................... 77-61 Monitor Engine Vibration .................................... 77-63 Fan Hub Balance Weights .................................. 77-65 Review Vibration Data ....................................... 77-67 PW1500G Trim Balance Helper Tool ................. 77-69
77-ii
CS130_M5_77_EngineIndicatingLOF.fm
77 - Engine Indicating
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77-iii
CS130_M5_77_EngineIndicatingLOF.fm
ENGINE INDICATING - CHAPTER BREAKDOWN ENGINE INDICATING CHAPTER BREAKDOWN
Speed (Power)
1
Temperature Indications
2
Vibration Monitoring
3
77 - Engine Indicating 77-11 Engine Power Indicating System
77-11 ENGINE POWER INDICATING SYSTEM GENERAL DESCRIPTION The engine power indicating system uses engine speed as a reference for engine power. The engine has three speed sensors that monitor engine speed: •
Fan speed probe
•
N1 speed probe
•
N2 speed sensor
The fan speed probe reads engine fan speed from the fan reduction gearbox front shaft gear. The N1 speed probe measures low-pressure compressor (LPC) speed. The N1 gauge has analog and digital readouts. N1 is the primary thrust control parameter. The thrust reference, calculated as an N1 value, is displayed by a reference bug on the analog display, and in digital format above the digital readout box. The N2 speed sensor measures high-pressure compressor (HPC) speed. The N2 speed sensor is part of the permanent magnet alternator generator (PMAG). The N2 is displayed in digital readout format only. The electronic engine control (EEC) receives the speed sensor signals. The EEC uses the signals for engine control and outputs indication signals for the N1 and N2 engine indication and crew alerting system (EICAS) display. The N1 gauges and the N2 digital readouts for both engines are displayed on EICAS. The fan speed is not displayed, but is used as backup in case of N1 signal failure.
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77-2
CS130_77_L2_EPIS_General.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
2 N1 SPEED PROBE
EEC
1 FAN SPEED PROBE
CLB 1
2 3
92.2
92.2
73.0
73.3
N1
718
722
EGT
86.5 1090 124 116
LEGEND
EEC PMAG
N1 (LPT and LPC) N2 (HPT and HPC) Nf (fan rotor) Electronic Engine Control Permanent Magnet Alternator Generator Analog ARINC 429
N2 FF (KPH) OIL TEMP OIL PRESS EICAS
86.6 1110 124 116
CS1_CS3_7700_005
3 PMAG
Figure 1: Engine Power Indicating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-3
CS130_77_L2_EPIS_General.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
COMPONENT LOCATION The engine power indicating system consists of the following components: •
Fan speed probe
•
N1 speed probe
•
Permanent magnet alternator generator (PMAG)
NF SPEED SENSOR The fan speed probe is located at the 1 o’clock position in the no.1 bearing support housing of the engine fan. N1 SPEED PROBE The N1 speed probe is mounted at station 2.5, at the rear of the compressor intermediate case at the 4:30 position. PERMANENT MAGNET ALTERNATOR GENERATOR The N2 speed coil is integral to the PMAG, which is located on the aft side of the main gearbox at the 7 o’clock position.
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77-4
CS130_77_L2_EPIS_ComponentLocation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
A
B
C
Fan Speed Probe
Permanent Magnet Alternator Generator
B N1 SPEED PROBE
C MAIN GEARBOX
CS1 CS3 7711 001 CS1_CS3_7711_001
A TOP VIEW
Figure 2: Engine Power Indicating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-5
CS130_77_L2_EPIS_ComponentLocation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
COMPONENT INFORMATION N1 SPEED PROBE The N1 speed probe is a dual sensor with two independent, isolated coils, and electrical connectors. The probe provides dual signals proportional to the rotational speed of the no. 2 bearing coupling nut located on the LPC. The EEC uses one signal per channel to generate the N1 value.
NOTE The N1 probe is approximately 20 cm (8 in.) in length. Care should be taken during removal and installation to avoid damaging the tip.
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77-6
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
B
Sensor A
A N1 SPEED PROBE
B COUPLING NUT
CS1_CS3_7711_003
Compressor Intermediate Case
Figure 3: N1 Speed Probe Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-7
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
PERMANENT MAGNET ALTERNATOR GENERATOR The PMAG stator has two speed coils that transmit two independent high rotor (N2) speed signals to the EEC. The PMAG stator also has dedicated coils for both EEC channel power generation, and fly-by-wire (FBW) power. The PMAG has offset captive bolts and transfer tubes to ensure proper clocking upon installation.
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77-8
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
Cover
Rotor
PERMANENT MAGNET ALTERNATOR GENERATOR STATOR
LEGEND EEC Channel A Power EEC Channel B Power Fly-By-Wire Power Speed (N2) Coils
CS1_CS3_7711_002
MAIN GEARBOX
Figure 4: Permanent Magnet Alternator Generator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-9
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
CONTROLS AND INDICATIONS N1 INDICATION The N1 indication consists of an analog and digital readout. A low-pressure compressor (LPC) speed of 100% N1 is equivalent to 10,600 rpm. The analog gauge consists of a pointer in a round arc, representing the operating range of the engine. The area below the pointer is shaded to provide a quick visual reference. At the upper end of the arc is a red line limit, indicating 100% N1 speed. The N1 digital display is in a white box to the right of the analog gauge. If the N1 speed is greater than 100%, the pointer moves above the red line limit. The pointer and the shaded area below changes to red. The arc extends above the red line and remains there even if the overspeed condition no longer exists. This indicates that the engine speed has exceeded 100%. The digital display and box color changes to red when N1 is above 100%. The display and box changes to white if the overspeed condition no longer exists. The N1 thrust reference bug readout is located on the outside of the analog N1 gauge arc.
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77-10
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
N1 REDLINE
Normal red line limit Overspeed Condition N1 overspeed condition
CLB N1 Red Line Limit
92.2
92.2
73.0
104
Takeoff Reference Without FLEX or MAX Climb
N1 N1 DIGITAL PARAMETER Symbol
EICAS
Description
86.5
Engine operation in normal range.
102
Engine operation above red line.
èèè
Invalid
CS1_CS3_7711_005
N1 Analog Pointer N1 Numeric Indication
Figure 5: N1 Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77-11
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
N1 THRUST REFERENCE BUG Other indications on the N1 display include an N1 thrust reference bug. The thrust reference bug is the best N1 for the selected thrust reference mode. The bug is displayed in magenta if the flight management system (FMS) calculates the thrust reference, or is displayed in cyan if it is entered manually into the FMS. N1 THRUST REFERENCE The N1 thrust reference is shown above the N1 digital display box. This value is displayed in magenta when calculated by the FMS, and in cyan when entered manually into the FMS. N1 THRUST COMMAND CURSOR The N1 thrust command cursor indicates the difference between the pointer and predicted N1 speed after the engine throttle is moved. When in automatic power reserve (APR) mode, the thrust reference bug changes to green. An offset remains in steady-state since the actual N1 is higher due to APR. Thrust Limitation at Low Speed The upper end of the thrust command cursor is shaded gray when thrust limitation at low speed is enabled during takeoff. The upper end of the gray area indicates the commanded N1, while the lower end indicates the thrust limitation. As the airspeed increases the gray area reduces and is removed when airspeed is above 80 kt. TAKEOFF REFERENCE WITHOUT FLEX OR MAX CLIMB The white radial line shows the thrust reference if FLEX was canceled. If FLEX was not used, the radial line is masked by the pointer. In climb, the white radial line shows the maximum climb thrust reference. If the information is invalid, the line is removed.
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77-12
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
Takeoff Reference if FLEX Was Canceled or MAX Climb Thrust Reference N1 Thrust Reference Bug
N1 THRUST REFERENCE BUG
N1 Thrust Reference
Symbol
Condition
Target N1 value entered manually in N1 data entry field
92.2
73.0
Bug position indicates target N1 value as computed automatically by the FMS
Thrust Command Cursor N1 SPEED N1 THRUST COMMAND CURSOR Symbol
Condition
Commanded decrease in N1
90.9
57.3 Commanded increase in N1 Thrust Limitation at Low Speed
APR APR Active
CS1_CS3_7711_006
COMMANDED INCREASE IN N1
Figure 6: N1 Thrust Reference Bug, N1 Thrust Reference, N1 Thrust Command Cursor, and Takeoff Reference Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-13
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
N2 INDICATION The N2 indication is a three-digit value located in the EICAS secondary engine indications area. The three-digit value represents a percentage of the N2 speed. In normal operation the digits are displayed in white. When N2 speed exceeds 100% the color changes to red.
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77-14
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
N2
86.6 1110 124 116
Symbol
Description
86.5
Engine operation in normal range
103.5
N2 above N2 red line threshold
èèè
Invalid
CS1_CS3_7711_007
EICAS
Figure 7: N2 Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-15
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
DETAILED DESCRIPTION Each channel of the electronic engine control (EEC) receives the following inputs: •
Nf speed
•
N1 speed
•
N2 speed
As long as the EEC is receiving one of the N1 speed signals, the EEC uses N1 as the primary thrust reference. In the event that one EEC channel fails, the other channel maintains engine control.
Data concentrator unit module cabinet 1 (DMC 1) and DMC 2 also receive calculated N1 and N2 information via an ARINC 429 BUS for distribution to the aircraft systems. A L (R) ENG EXCEEDANCE warning message indicates overspeed of the N1 or N2 speeds. If the EEC detects a N1 or N2 speed in excess of 105%, the EEC initiates an automatic shutdown of the engine. In the event the N2 value drops below 5% of the engine governed speed, an L(R) ENG FAIL caution message is displayed.
In the event of a loss of the both N1 speed signals, the EEC calculates the N1 speed using the Nf signal. In the case of failure of one of the dual N1 speed signals, a L (R) ENGINE FAULT advisory message indicates a fault affecting the dispatch of the aircraft. The EEC uses the N2 speed signal for control of the fuel and ignition during starting and overspeed monitoring. In the event that one EEC channel fails, the other channel provides N2 speed. In the event of a loss of one of the N2 signals, a L(R) ENG OPER DEGRADED caution message is displayed on the engine indication and crew alerting system (EICAS). A L(R) ENG FAULT - FADEC FAULT 1 crew alerting system (CAS) advisory message is also displayed and rapid thrust movement is to be avoided. If both N2 signals are lost, the EEC commands an engine shutdown and the L(R) ENG FAIL caution CAS message is displayed. The EEC sends the N1, N2, and Nf information to the prognostics and health management unit (PHMU) via CAN DATA BUS. The PHMU uses the N1 and N2 speeds for engine vibration monitoring. The Nf is used by the PHMU for engine fan trim balancing. The EEC sends the calculated N1 and N2 information directly to the EICAS via an ARINC 429 BUS. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-16
CS130_77_L3_EPIS_DetailedDescription.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
CLB FAN SPEED PROBE Prognostics and Health Management Unit
Nf Speed
92.2
92.2
73.0
73.3
N1 EEC
N1 SPEED PROBE
N1 Speed 1 N1 Speed 2
718
Channel A
722
EGT PMAG
Channel B N2 Speed 1 N2 Speed 2
DMC 1 and DMC 2
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
LEGEND
EICAS
Data Concentrator Unit Module Cabinet Electronic Engine Control Permanent Magnet Alternator Generator Analog ARINC 429 CAN BUS
NOTE Left engine shown. Right engine similar.
CS1_CS3_7711_004
DMC EEC PMAG
86.6 1110 124 116
Figure 8: Engine Power Indicating System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-17
CS130_77_L3_EPIS_DetailedDescription.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and fault messages for the engine power indicating system.
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77-18
CS130_77_L2_EPIS_MonitoringTests.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
CAS MESSAGES
Table 3: FAULT Messages Table 1: WARNING Messages
MESSAGE DUAL ENG FAIL
MESSAGE
LOGIC L ENG and R ENG switches are ON and dual engine flameout.
LOGIC
L(R) ENG FAULT- INFO message will be set when any non-dispatchable FADEC FAULT 1 fault has been detected by the left (right) engine FADEC system or the DMC receives no valid data from left (right) engine for 10 seconds.
Table 2: CAUTION Messages MESSAGE
LOGIC
L ENG EXCEEDANCE
L N1 or L N2 or L ITT or L oil temperature above threshold.
R ENG EXCEEDANCE
R N1 or R N2 or R ITT or R oil temperature above threshold.
L ENG FAIL
L ENG switch in ON with the engine in a sub-idle condition and a relight or restart was unsuccessful.
R ENG FAIL
R ENG switch ON with the engine in a sub-idle condition and a relight or restart was unsuccessful.
L ENG OPER DEGRADED
Left EEC has detected a loss of thrust or degraded thrust control due to:
R ENG OPER DEGRADED
•
Fuel flow or actuator not tracking to commanded position
•
Input parameter in failsafe
•
Engine flameout
Right EEC has detected a loss of thrust or degraded thrust control due to: •
Fuel flow or actuator not tracking to commanded position
•
Input parameter in failsafe
•
Engine flameout
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77-19
CS130_77_L2_EPIS_MonitoringTests.fm
77 - Engine Indicating 77-21 Temperature Indicating System
77-21 TEMPERATURE INDICATING SYSTEM GENERAL DESCRIPTION The temperature indicating system measures the exhaust gas temperature (EGT) for display on the engine indication and crew alerting system (EICAS). EGT is monitored at station 5 of the engine. Two EGT probe and cable assemblies are used to transmit the EGT signals to the main oil temperature (MOT) probe and then to the electronic engine control (EEC). The oil temperature probe is used as the cold junction for the EGT signal. The EEC then calculates the EGT value and sends the information to the EICAS display. Each EGT probe and cable assembly consists of four EGT probes. The assemblies are semi-rigid and can only be replaced as an assembly. Individual EGT probes are not replaceable.
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77-20
CS130_77_L2_TIS_General.fm
77 - Engine Indicating 77-21 Temperature Indicating System
CLB Left EGT Probe and Cable Assembly
92.2
92.2
73.0
73.3
N1
OIL TEMPERATURE PROBE
718
ELECTRONIC ENGINE CONTROL
EGT
86.5 1090 124 116
Right EGT Probe and Cable Assembly
722
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
CS1_CS3_7721_001
EICAS
LEGEND EGT
Exhaust Gas Temperature ARINC 429
Figure 9: Exhaust Gas Temperature Indicating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-21
CS130_77_L2_TIS_General.fm
77 - Engine Indicating 77-21 Temperature Indicating System
COMPONENT LOCATION The temperature indicating system consists of the following components: •
Exhaust gas temperature (EGT) probe and cable assembly
•
Oil temperature probe
•
Probe junction
EXHAUST GAS TEMPERATURE PROBE AND CABLE ASSEMBLY The EGT probe and cable assemblies are mounted around the circumference of the turbine exhaust case (TEC). There is one assembly on the left side of the engine, and one on the right side. OIL TEMPERATURE PROBE The oil temperature probe is installed on the oil control module (OCM) located on the right side of the engine. PROBE JUNCTION The probe junction is located on each side of the TEC at the 3 o’clock and 9 o’clock position.
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77 - Engine Indicating 77-21 Temperature Indicating System
A
B
EGT Probe and Cable Assembly
B PROBE JUNCTION
OIL TEMPERATURE PROBE
CS1_CS3_7721_006
A
Figure 10: Exhaust Gas Temperature Indicating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77-23
CS130_77_L2_TIS_ComponentLocation.fm
77 - Engine Indicating 77-21 Temperature Indicating System
CONTROLS AND INDICATIONS INDICATIONS The exhaust gas temperature (EGT) display shows the EGT for each engine on the EICAS display, as calculated by the EEC. The gauges consist of an analogue and a four-digit display. The display is shown in degrees Celsius (°C). The analogue gauge consists of a pointer in a round arc representing the operating range of the engine. The area below the pointer is shaded to provide a quick visual reference. At the upper end of the arc is a red line limit, indicating maximum operating temperature. There is also a yellow line to indicate approaching maximum temperature limit. The electronic engine control (EEC) calculates an engine operating time limit in the yellow range. If the time limit exceeds the EGT needle, the shaded area and digital readout change to red. The yellow line is removed during takeoff, go-around, reverse, or when invalid. When the EGT pointer goes above the yellow line, the pointer and the shaded area below the pointer change to yellow. When the EGT pointer goes above the red line limit, the pointer and the shaded area below the pointer change to red. The arc is also extended in red above the red line limit, and stays there even if the overtemperature condition goes away. This indicates that the engine temperature has exceeded its limit. The digital EGT display changes color to match the pointer and shaded area.
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77 - Engine Indicating 77-21 Temperature Indicating System
EGT INDICATIONS
860 718
Engine start on ground only Yellow limit removed after engine start completed
722
EGT
1036
EICAS
EGT above amber line Engine start on ground only
Normal Indications - Engine Running on ground after start and in flight
1064
CS1_CS3_7721_005
EGT above red line
Figure 11: Exhaust Gas Temperature Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-25
CS130_77_L2_TIS_ControlsIndications.fm
77 - Engine Indicating 77-21 Temperature Indicating System
DETAILED DESCRIPTION The four signals from both EGT probe and cable assemblies are electrically averaged into two signals, which are sent to the oil temperature probe for compensation. The oil temperature probe sends compensated EGT signals to the EEC. The left EGT probe and cable assembly signal is sent to channel A. The signal of the right EGT probe and cable assembly is sent to channel B of the EEC. The EEC converts the average EGT analog signal to a digital signal. The two signals are averaged into a single output, which is sent directly to the display unit (DU) for display on the EICAS. A L(R) ENGINE FAULT advisory message is displayed when one of the EGT probe and cable assemblies fails.
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77 - Engine Indicating 77-21 Temperature Indicating System
Left EGT Probe and Cable Assembly
OIL TEMPERATURE PROBE
ELECTRONIC ENGINE CONTROL
Probe Junction Channel A
Channel A
Channel B
Channel B
718
Probe Junction
EGT EICAS
Right EGT Probe and Cable Assembly
EGT
CS1_CS3_7721_004
LEGEND Exhaust Gas Temperature ARINC 429 CAN BUS
Figure 12: Exhaust Gas Temperature Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-27
CS130_77_L3_TIS_DetailedDescription.fm
77 - Engine Indicating 77-21 Temperature Indicating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the engine temperature indicating system.
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77 - Engine Indicating 77-21 Temperature Indicating System
CAS MESSAGES Table 4: CAUTION Messages MESSAGE
LOGIC
L ENG EXCEEDANCE
L N1 or L N2 or L ITT or L oil temperature above threshold.
R ENG EXCEEDANCE
R N1 or R N2 or R ITT or R oil temperature above threshold.
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77 - Engine Indicating 77-32 Vibration Monitoring
77-32 VIBRATION MONITORING GENERAL DESCRIPTION The engine vibration indicating system monitors, records, and analyzes engine vibration before displaying this information on the engine indication and crew alerting system (EICAS). Engine speed sensors are used, in conjunction with vibration monitors, to help determine which rotor may be out of balance. The prognostics and health management unit (PHMU) compares engine speed data from the electronic engine control (EEC) with vibration data sent from the forward and aft rotor vibration sensors. The PHMU receives power from DC BUS 1. Data stored in the EECs data storage unit (DSU) over a period of flights can be used by the onboard maintenance system (OMS) to conduct engine vibration surveys. Vibration sensors are used on both N1 and N2 rotors to send signals that are proportional to engine vibration. The PHMU converts the signals from charge to voltage, filters them, and correlates the frequencies with the specific rotor. The N1 and N2 vibration signals are then sent to the EEC for display on EICAS. These signals are also used by the aircraft OMS to calculate solutions for correcting a fan imbalance.
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77 - Engine Indicating 77-32 Vibration Monitoring
CLB DC BUS 1
92.2
92.2
73.0
73.3
N1
DSU
718
722
EGT
Speed Sensing FWD VIBRATION SENSOR
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
EICAS PROGNOSTICS AND HEALTH MANAGEMENT UNIT
ELECTRONIC ENGINE CONTROL
MAINT MENU EEC1-A-TRIM BALANCE PROCEDURE MAIN MENU 4/90
AFT VIBRATION SENSOR REVIEW VIB DATA HISTORY
REVIEW/EDIT STORED WEIGHTS
PERFORM BALANCE
DISPLAY BALANCE COEFFICIENTS
LEGEND DSU
MODIFY SETTINGS
Data Storage Unit Analog ARINC 429
Write Maintenance Reports
OMS
Cancel
CS1_CS3_7732_004
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
Figure 13: Vibration Monitoring Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_77_L2_VM_General.fm
77 - Engine Indicating 77-32 Vibration Monitoring
COMPONENT LOCATION The vibration monitoring system consists of the following components: •
Forward vibration sensor
•
Aft vibration sensor
•
Prognostics health and management unit (PHMU)
•
Data storage unit (DSU)
FORWARD VIBRATION SENSOR The forward vibration sensor is mounted on the compressor intermediate case at the 9 o’clock position. AFT VIBRATION SENSOR The aft vibration sensor is mounted on the low-pressure turbine (LPT) housing at the 3 o’clock position. PROGNOSTICS HEALTH AND MANAGEMENT UNIT The prognostics health and management unit (PHMU) is mounted to the fan case adjacent to the EEC. DATA STORAGE UNIT The data storage unit (DSU) is located on the EEC.
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77 - Engine Indicating 77-32 Vibration Monitoring
A
A AFT VIBRATION SENSOR RIGHT SIDE VIEW Electronic Engine Control
Data Storage Unit
Prognostics and Health Management Unit
LEFT SIDE VIEW
B
FWD VIBRATION SENSOR
CS1_CS3_7731_001
B
Figure 14: Vibration Monitoring Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-33
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77 - Engine Indicating 77-32 Vibration Monitoring
CONTROLS AND INDICATIONS INDICATIONS The following vibration indications are displayed in the EICAS: •
N1 vibration is displayed as an AMBER VIB flag in the associated N1 gauge
•
N2 vibration is displayed as an AMBER VIB flag adjacent to the digital readout
•
FAN VIB is displayed in units
The N1 VIB is displayed as a flag in the N1 gauge when the parameter exceeds the normal vibration threshold of 0.52 inches per second (IPS). The REV icon replaces the N1 VIB display when the thrust reversers deploy. The N2 VIB is displayed as a flag adjacent to the N2 speed when the parameter exceeds the normal vibration threshold of 1.2 IPS. The FAN VIB level is posted on EICAS in units from 0 to 8. The units represent a percentage of the vibration threshold where 4.0 is 100%. The normal vibration threshold for the engine fan is 1.2 IPS. FAN VIB is only displayed when an engine fan vibration exceeds 4.0 units, or a fault has occurred in the vibration monitoring system. When this occurs, both left and right engine FAN VIB is displayed. The fan vibration exceedance is displayed in amber, while normal indications are displayed in white.
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77 - Engine Indicating 77-32 Vibration Monitoring
CLB 92.2
92.2
73.0
73.3
FAN VIBRATION Symbol
N1
718
VIB
722
0.6 4.2 èèè
Description Vibration at or below high vibration threshold Vibration above high vibration threshold If both vibration monitors fail on one side.
EGT N1 and N2 VIBRATION
86.5 VIB N2 86.6 1090 FF (KPH) 1110 124 OIL TEMP 124 116 OIL PRESS 116 FAN VIB 4.2 0.4
Symbol
VIB
Description Vibration above high vibration threshold N1 VIB is not displayed when the thrust reverser is operating.
CS1_CS3_7732_001
EICAS
Figure 15: Vibration Monitoring Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
DETAILED DESCRIPTION Vibration monitoring is performed through data input from the two engine vibration sensors, as well as speed inputs from the EEC for N1, N2, and the fan speed (Nf). The PHMU receives engine speed signals from the EEC and vibration signals from the vibration sensors. The Nf speed signal is used in conjunction with the forward vibration sensor to measure fan vibration. The PHMU compares the engine speeds and vibration signals to determine the vibration level for the fan, N1, and N2 shafts. The PHMU sends the vibration signals to the EEC via a CAN BUS. The EEC sends vibration information for display on EICAS. When engine vibration exceedances are detected, an ENG VIBRATION caution message is displayed, indicating excessive vibration levels. Engine vibration information is also sent to the DMCs for the OMS and health management unit (HMU).
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77 - Engine Indicating 77-32 Vibration Monitoring
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
CDC 1-10-8 2 PHMU DC BUS 1
SSPC 5A
Logic: Always on
ELECTRONIC ENGINE CONTROL N1 Signal N2 Signal
Channel A
Channel B
EICAS
HMU
DMC 1 and DMC 2
OMS
Nf Signal
FWD VIBRATION SENSOR
AFT VIBRATION SENSOR
DMC
Data Concentrator Unit Module Cabinet ARINC 429 CAN BUS NOTE Left engine shown.
CS1_CS3_7731_002
LEGEND
Figure 16: Vibration Monitoring Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine vibration monitoring system.
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77 - Engine Indicating 77-32 Vibration Monitoring
CAS MESSAGES Table 5: CAUTION Messages MESSAGE ENGINE VIBRATION
LOGIC Left or right engine vibration (either fan vibration, N1 or N2).
Table 6: ADVISORY Message MESSAGE
LOGIC
L (R) ENG FAULT Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 7: INFO Messages MESSAGE
LOGIC
77 L (R) ENGINE FAULT - PHMU INOP
FADEC logic has detected that the PHMU is failed.
77 L (R) ENGINE FAULT - N1/FAN VIBRATION NON DEGRADED
INFO message will be set when a forward vibration sensor fault has been detected or N1 or NF speed input for vibration processing has failed.
77 L (R) ENGINE FAULT - N2 VIBRATION NON DEGRADED
INFO message will be set when a rear vibration sensor fault has been detected or N2 speed input for vibration processing has failed.
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77 - Engine Indicating 77-32 Vibration Monitoring
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM Trim Balance The purpose of the trim balance procedure is to provide the information needed to balance the fan and/or low spool of the engine. The fan is a large rotating mass, even minor blade damage can cause an imbalance. A trim balance procedure is used to reduce fan-related vibration onwing. While the engine is running, the prognostics and health management unit (PHMU) continuously collects data for trim balancing using values from the vibration sensors and the predetermined Nf, N1, and N2 engine speeds. Historical vibration data stored in the EECs data storage unit (DSU) is used as reference in calculating the trim balance solutions. The system corrects small amounts of residual imbalance, which could be related to minor blade damage or fan blade replacement, even when matched sets are replaced.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU EEC1-A-TRIM BALANCE PROCEDURE MAIN MENU 4/90
REVIEW VIB DATA HISTORY Data Storage Unit
REVIEW/EDIT STORED WEIGHTS DMC
PERFORM BALANCE
DISPLAY BALANCE COEFFICIENTS ELECTRONIC ENGINE CONTROL
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT PROGNOSTICS AND HEALTH MANAGEMENT UNIT
LEGEND DMC
Write Maintenance Reports
Data Concentrator Unit Module Cabinet
ONBOARD MAINTENANCE SYSTEM
Cancel
CS1_CS3_7730_001
MODIFY SETTINGS
Figure 17: Trim Balance Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-41
CS130_77_L2_VM_PracticalAspects.fm
77 - Engine Indicating 77-32 Vibration Monitoring
Fan Hub Balance Weights N1 (fan rotor) vibration signals are sent to the PHMU and the EEC for interpretation and calculation of optimal trim balance solutions. The onboard maintenance system (OMS) uses these solutions to indicate where to install balance weights on the fan rotor disc. Each weight has a different part number, based on its specific weight. The fan speed sensor is used to determine the fan index position sensing. It measures the fan index rotational position, once-per-revolution, to allow the location of vibration imbalances to support the fan trim balance. The trim balance procedure involves running a baseline (as-is) vibration survey. The final correction weight is then installed on a balance weight flange, located on the front of the fan rotor hub. The software-stored weight configuration should exactly match the actual balance weight configuration on the engine in order for the OMS system-provided balance solution to be correct. Reviewing and editing the stored balance weight configuration is a precondition for performing a trim balance. A 0° reference point is aligned with a dimple on the fan hub shaft. Trim balance weights are installed on either side of the module balance weights. Module balance weights are riveted in place during production or engine rebuild, and should not be disturbed during trim balance procedures.
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77 - Engine Indicating 77-32 Vibration Monitoring
0° Reference for Weight Angles Aligned With a Shaft Dimple
Fan Drive Shaft Dimple Location
Trim Balance Locations (for clip-on weights)
VIEW FORWARD LOOKING AFT (fan drive shaft shown)
CS1_CS3_7730_009
Z-Balance and Module Balance Locations (for riveted weights)
Figure 18: Fan Hub Balance Weights Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Trim Balance Procedure Main Menu On the OMS maintenance page, the trim balance procedure main menu screen provides an entry point into the trim balance subprocedures, including: •
REVIEW VIB DATA HISTORY
•
REVIEW/EDIT STORED WEIGHTS
•
PERFORM BALANCE
•
DISPLAY BALANCE COEFFICIENTS
•
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
•
MODIFY SETTINGS
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU EEC1-A-TRIM BALANCE PROCEDURE MAIN MENU 4/90
REVIEW VIB DATA HISTORY
REVIEW/EDIT STORED WEIGHTS BEFORE A BALANCE SOLUTION IS IMPLEMENTED THE CURRENT WEIGHT CONFIGURATION STATUS SHOULD BE REVIEWED AND CORRECTED IF NEEDED.
PERFORM BALANCE BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM LAST COMPLETED RUN OR BASED ON SEVERAL PAST RUNS.
DISPLAY BALANCE COEFFICIENTS
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
MODIFY SETTINGS
Write Maintenance Reports
Cancel
ONBOARD MAINTENANCE SYSTEM
CS1_CS3_7730_002
THIS PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY BE USED IF THE GENERIC COEFFICIENTS DO NOT PRODUCE SATISFACTORY RESULTS.
Figure 19: Trim Balance Procedure Main Menu Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Review VIB Data History REVIEW VIB DATA HISTORY determines changes to the overall vibration levels in the recent flight history.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
EEC1-A-VIB HISTORY MENU 4/90 0
10/90 -----------TRIM BALANCE RELATED VIB HISTORY-----------
REVIEW VIB DATA HISTORY
FAN VIB HISTORY
REVIEW/EDIT STORED WEIGHTS GHTS LPC VIB HISTORY
LPT VIB HISTORY
PERFORM BALA BALANCE BALANC ALAN NCE E
----------OTHER BALANCE RELATED VIB HISTORY----------
HP SPOOL VIB HISTORY DISPLAYY BALANCE COEFFICIENTS BROADBAND VIB HISTORY CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT WEIGHT ----VIEW A SUMMARY OF AVAILABLE BALANCE SOLUTIONS----
MODIFY SETTINGS GS
RETURN TO TRIM BALANCE MAIN MENU
Write Maintenance M Reports ports ts
Cance Cancel ncel l
Write Maintenance Reports
Cancel
CS1_CS3_7730_003
VIEW BALANCE SOLUTION SUMMARY
Figure 20: Review VIB Data History Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Review/Edit Stored Weights REVIEW/EDIT STORED WEIGHTS provides a starting point for examining and changing software stored balance weight configurations for each of the correction planes (fan, LPC, and LPT).
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
GPS 1-REVIEW/EDIT STORED WEIGHT CONFIGURATION 4/90 0 THIS PROCEDURE WILL REQUIRE ACCESS TO BALANCE CORRECTION PLANES TO EXAMINE CURRENT STATE.
REVIEW VIB DATA HISTOR HISTORY TORY Y
REVIEW/EDIT STORED WEIGHTS
REVIEW/EDIT FAN CORRECTION PLANE WEIGHTS
PERFORM BALA BALANCE BALANC ALAN NCE E
REVIEW/EDIT LPC CORRECTION PLANE WEIGHTS DISPLAYY BALANCE COEFFICIENTS REVIEW/EDIT LPT CORRECTION PLANE WEIGHTS
MODIFY SETTINGS GS
RETURN TO TRIM BALANCE MAIN MENU
Write Maintenance M Reports ports ts
Cance Cancel ncel l
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7730_004
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT WEIGHT
Figure 21: Review/Edit Stored Weights Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-49
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77 - Engine Indicating 77-32 Vibration Monitoring
Perform Balance PERFORM BALANCE provides options for selecting which type of balance procedure to perform. Each of the provided buttons lead to separate balance solutions, based on the selected balance plane (fan, LPC or LPT). All balance solutions are precalculated and stored during previous engine runs.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
EEC1-A-SELECT BALANCE PROCEDURE 4/90 0
4/90 THIS PROCEDURE WILL REQUIRE ACCESS TO BALANCE DO PLANES FAN TRIM BALANCECURRENT (SINGLE PLANE) CORRECTION TO EXAMINE STATE.
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y
PROCEDURE WILL REQUIRE FAN CORRECTION PLANE ACCESS.
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
DO LPT TRIM BALANCE (SINGLE PLANE)
PERFORM BALANCE OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM RO LAST ASSTT
PROCEDURE WILL REQUIRE LPT CORRECTION PLANE ACCESS.
DISPLAY BALANCE COEFFICIENTS IENTS
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT CA
MODIFY DIFY SETTINGS SETTINGS
Write Maintenance M Reports Repor ports ts
RETURN TO TRIM BALANCE MAIN MENU
Cance Cancel ancel l
Write Maintenance Reports
Cancel
CS1_CS3_7730_005
S PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY THIS S ICC COEFFICIENTS RODUCE BE USED PRODUCE IF THE GENERIC CCO OEFFI OOEFFICIENTS IC CI CI NNTSS DO DO NO NOT OT PRO PRODUC PPRODU ROD O UC CE SATISFACTORY RESULTS. .
Figure 22: Perform Balance Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-51
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77 - Engine Indicating 77-32 Vibration Monitoring
Display Balance Coefficients The DISPLAY BALANCE COEFFICIENTS page is an advanced screen used to modify OMS trim balance settings. Refer to the Aircraft Maintenance Publication (AMP) for guidance as to when this advanced subprocedure should be used.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
GPS 1-REVIEW/EDIT STORED WEIGHT CONFIGURATION 4/90 0
8/90
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y FAN GENERIC NUM MAG PHASE 1: X.X XXX 2: X.X XXX 3: X.X XXX 4: X.X XXX 5: X.X XXX 6: X.X XXX 7: X.X XXX 8: X.X XXX 9: X.X XXX 10: X.X XXX
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
PERFORM BALANC BALANCE BALAN NCE E
NUM 11: 12: 13: 14: 15: 16: 17: 18: 19: 20:
MAG X.X X.X X.X X.X X.X X.X X.X X.X X.X X.X
PHASE XXX XXX XXX XXX XXX XXX XXX XXX XXX XXX
OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM ROM RO LAST ASSTT
EACH SET CONTAIN 20 PAIRS OF COEFFICIENTS EACH PAIR = A SENSITIVITY FACTOR AND A LAG ANGLE
DISPLAY BALANCE COEFFICIENTS
E-FLANGE 1...................10
EACH BLOCK OF 10 PAIRS ARE PER THE 10 SPEED POINTS
S PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY THIS S ICC COEFFICIENTS RODUCE BE USED IF THE GENERIC PRODUCE CCO OEFFI OOEFFICIENTS IC CI CI NNTSS DO DO NO NOT OT PRO PRODUC PPRODU ROD O UC CE SATISFACTORY RESULTS. .
RETURN TO SELECT COEFFICIENT SCREEN
MODIFY SETTINGS SETTING ETTINGS S
Write ite Maintenance M Reports ports orts
Cance Cancel ncel l
Write Maintenance Reports
Cancel
CS1_CS3_7730_006
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT CA
P-FLANGE 11..................20
Figure 23: Display Balance Coefficients Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Engine Serial Number Coefficients The calculated ESN coefficients, using the trial weight screen, should only be used if the normal procedure does not produce the desired reduction in vibration. The screen enables the installation of trial weights to help determine a solution.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
GPS 1-CALCULATE BALANCE COEFFICIENTS-1 4/90 0
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y
TO CALCULATE BALANCE COEFFICIENTS, THE OPERATOR MUST COMPLETE THE FOLLOWING STEPS IN ORDER 1) SELECT DESIRED BALANCE PLANE 2) REVIEW AND CORRECT CURRENT WEIGHT CONFIGURATION
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
3) ADD A TRIAL WEIGHT TO BALANCE PLANE 4) EXIT INTERACTIVE MODE TO CONDUCT A TRIM BALANCE SURVEY PER AMM
PERFORM BALANC BALANCE BALAN NCE E
5) RETURN TO TRIM BALANCE INTERACTIVE PROCEDURE TO CONFIRM REMOVAL OF TRIAL WEIGHT AND TO REVIEW NEW BALANCE COEFFICIENT
OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM ROM RO LAST ASSTT
DISPLAY BALANCE COEFFICIENTS OEFFICIENTS
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
SELECT "CONTINUE" TO START STEP 1...
MODIFY SETTINGS SETTING ETTINGS S
Write ite Maintenance M Reports ports orts
RETURN TO TRIM BALANCE MAIN MENU
Cance Cancel ncel l
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7730_007
THIS PROCEDURE RE REQUIRES ROUND RUN HOULD QUIRES A GRO GROUND RU AND SSHOULD OULD ONLY O S ICC COEFFICIENTS OEFFICIENTS S DO RODUCE BE USED PRODUCE IF THE GGENERIC CCO OOEFFI IC CI CI NNTS DO NO NOT OT PRO PRODUC PPRODU ROD O UC CE SATISFACTORY RESULTS. .
Figure 24: Engine Serial Number Coefficients Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Modify Settings Screen The MODIFY SETTINGS screen is an advanced screen that allows setting changes for trim balance processing, clearing configuration, and balance histories.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y
GPS 1-MODIFY TRIM BALANCE SETTINGS 1/90 * NOTE * TRIM BALANCE SETTING ARE USED TO CONFIGURE SOFTWARE PROCESSING WITH THE SELECTIONS PROVIDED BELOW.
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
CLEAR BALANCE VECTOR HISTORY
4/90 0
BALANCE DATA FROM PAST RUNS WILL NOT BE USED FOR FUTURE BALANCE SOLUTION CALCULATIONS
PERFORM BALANC BALANCE BALAN NCE E
EXCLUDE GROUND RUN FROM UPDATING HISTRY
OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM ROM RO LAST ASSTT
PREVENT NORMAL OPERATION GROUND RUNS FROM UPDATING BALANCE VECTOR HISTORY
DISPLAY BALANCE COEFFICIENTS OEFFICIENTS
CLEAR STORED VIB HISTORY
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT CA
CLEAR ALL STORED BALANCE WEIGHT CONFIGURATION
MODIFY SETTINGS
RETURN TO TRIM BALANCE MAIN MENU
Write Maintenance M Reports ports orts
Cance Cancel ncel l
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7730_008
S PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY THIS S ICC COEFFIC RODUCE COEFFICIENTS PRODUCE BE USED IF THE GENERIC CO COEFFICI O CIENTS IENTS NNTSS DO DO NOT NO PRO PPRODUC PRODU ROD O UC CE SATISFACTORY RESULTS. .
Figure 25: Modify Settings Screen Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
FAN TRIM BALANCE USING THE PW 1500G TRIM BALANCE HELPER Test Set Up The Pratt & Whitney Special Instruction no. 44F - 16 Rev. A provides instructions for performing an on-wing fan trim balance using the onboard maintenance system (OMS) and the PW1500G Fan Trim Balance Helper Tool. Use the OMS to get weight and angle data for the fan. Use this data with the PW1500G Trim Balance Helper Tool provided in this Special Instruction to develop a balance solution. Perform a baseline engine vibration survey as follows: 1. Start both engines and allow engines to idle for a minimum of 5 minutes to allow the engine to become thermally stable. Make sure all the engine parameters are within the allowable limits. 2. For the engine under test, slowly move the thrust lever forward to 67-69% N1 speed and let the engine become stable for 3 minutes. The engine not being tested can be operated as required during this survey. 3. For the tested engine, slowly move the thrust lever back to MINIMUM IDLE power and let the engine become stable for 3 minutes.
NOTE Do not run the tested engine at idle power for more than 3 minutes. The engine can cool and the thermal stabilization steps must be repeated.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77 - Engine Indicating 77-32 Vibration Monitoring
Data Storage Unit
DMC
ELECTRONIC ENGINE CONTROL
LEGEND DMC
Data Concentrator Unit Module Cabinet ONBOARD MAINTENANCE SYSTEM
CS1_CS3_7730_001
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
Figure 26: Fan Trim Balance Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Onboard Maintenance System Data Prior to performing a baseline vibration survey, access the onboard maintenance system (OMS) system as follows: 1. Select Perform LRU/Systems Operations 2. Select ATA 73 - Engine Fuel & Control 3. Select Left or Right Engine EEC for the engine under test 4. Select Vibration from the drop-down menu to monitor the engine vibration data At the completion of the vibration survey and prior to engine shutdown, Data Reader is selected to get the fan trim balance solution weight and solution angle.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU
MAINT MENU Maintenance Main Menu
View Post Flight Summary View Flight Deck Effects (5 Actives)
View System Exceedances (4
Stored)
LRU/System Operations
LRU/System Operations
LRU/System Operations
View
All ATAs
View
All ATAs
View
All ATAs
Select ATA
21-00-00 AIR CONDITIONING
Select ATA
73-00-00 ENGINE FUEL & CONTROL
Select ATA
73-00-00 ENGINE FUEL & CONTROL
Select LRU/System
Prev 38-00-00 44-00-00 46-00-00 47-00-00 49-00-00 52-00-00 73-00-00
Select LRU/System Left Engine EEC (1)
View Fault Messages (5 Actives) View Service Messages (0 Active)
MAINT MENU
DATA Data Reader
WATER/WASTE CABIN CONTROLLER INFORMATION SYSTEMS FUEL TANK INERTING APU-GENERAL DOORS ENGINE FUEL & CONTROL
DATA Actuators
Left Engine EEC (1) Left Engine EEC (1) channel B Right Engine EEC (2) Right Engine EEC (2) channel A Right Engine EEC (2) channel B
Select LRU/System Left Engine EEC (1)
DATA Actuators
Air Data
Air Data
Cowl Anti-Ice
Cowl Anti-Ice
Data Reader
Data Reader
Fuel System
Fuel System
Oil System
Oil System
Performance
Performance
Starting
Starting
Thrust Ratings
Thrust Ratings
Write Maintenance Reports
Thrust Reverser
Thrust Reverser
Utility Functions
Vibration
Vibration
View System Trends View Aircraft Life Cycle Data View System Parameters Perform LRU/System Operations
CS1_CS3_7730_010
View System Configuration Data
Figure 27: Onboard Maintenance System Data Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Monitor Engine Vibration The engine vibration data for N1, N2, and the fan Nf can be monitored during the engine run. For the engine being tested, slowly move the thrust lever to achieve the following settings: •
Perform a slow acceleration (40-45 sec) to 60 +/-1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 64 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 68 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 72 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 76 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 80 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 83 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow deceleration (60-90 sec) to MINIMUM IDLE
NOTE The maximum speed may be limited by TLA rating. The ENG VIBRATION caution CAS can be disregarded for 30 seconds during this test in order to obtain vibration data as long as it is related to fan vibration.
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TECHNICAL TRAINING MANUAL
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU LRU/System Operations View
All ATAs
Select ATA
73-00-00 ENGINE FUEL & CONTROL
Select LRU/System Left Engine EEC (1)
DATA Actuators Air Data Cowl Anti-Ice Data Reader Fuel System Oil System Performance Starting Thrust Ratings Thrust Reverser
CS1_CS3_7730_012
Vibration
Figure 28: Monitor Engine Vibration Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
77-63
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77 - Engine Indicating 77-32 Vibration Monitoring
Fan Hub Balance Weights A 0° reference point is aligned with a dimple on the fan hub shaft. Trim balance weights are installed on either side of the module balance weights. Module balance weights are riveted in place during production or engine rebuild, and should not be disturbed during trim balance procedures. The fan hub balance weights can be accessed as follows: 1. Remove the inlet cone cover from the inlet cone. 2. Rotate the reference dimple marked on the fan drive shaft to the top dead center (TDC) position (12:00). 3. Note and record any previously installed trim balance weights (spring loaded weights only) and hole location.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77 - Engine Indicating 77-32 Vibration Monitoring
0° Reference for Weight Angles Aligned With a Shaft Dimple
Fan Drive Shaft Dimple Location
Trim Balance Locations (for clip-on weights)
VIEW FORWARD LOOKING AFT (fan drive shaft shown)
CS1_CS3_7730_009
Z-Balance and Module Balance Locations (for riveted weights)
Figure 29: Fan Hub Balance Weights Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Review Vibration Data Access the Data Reader page to get the following data: •
Record the values displayed for the following parameters: -
Label 47 Fan Trim Balance Solution Weight
-
Label 54 Fan Trim Balance Solution Angle
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT AINT MENU MENU
MAINT MENU
L EEC - Data Reader
Select ATA
All ATAs
Refre Refresh freshh
73-00-00 ENGINE FUEL & CONTROL
- LRU: Left Engine EEC (1) + Label: 43 N1 Vibration - Front Vib Senso + Label: 43 N1 Vibration - Front Vib Senso + Label: 44 N2 Vibration - Rear Vib Sensor + Label: 44 NF Vibration - Rear Vib Sensor + Label: 45 NF Vibration - Rear Vib Sensor + Label: 45 NF Vibration - Rear Vib Sensor + Label: 46 N1 Phase Angle - Front Vib Sen + Label: 47 N1 Phase Angle - Front Vib Sen - Label: 47 Fan Trim Balance Solution Weig SDI 1 Fan Trim Balance Solution Weight 0.0000 SSN [BNR] 3 - Label: 47 Fan Trim Balance Solution Weig SDI 1 Fan Trim Balance Solution Weight 0.0000 SSN [BNR] 3 - Label: 54 Fan Trim Balance Solution Angl SDI 1 Fan Trim Balance Solution Angle_ 0.0000 SSN [BNR] 3 - Label: 54 Fan Trim Balance Solution Angl SDI 3 Fan Trim Balance Solution Angle_ 0.0000 SSN [BNR] 3 + Label: 55 LPT Trim Balance Solution Weig + Label: 55 LPT Trim Balance Solution Weig + Label: 56 LPT Trim Balance Solution Angl + Label: 56 LPT Trim Balance Solution Angl + Label: 60 ODM Sensor Raw Signal RMS Valu
Select LRU/System Left Engine EEC (1)
DATA Actuators Air Data Cowl Anti-Ice Data Reader Fuel System Oil System Performance Starting Thrust Ratings Thrust Reverser Vibration
Write Mai Maintenance ce Reports orts
+ All
- AAll ll Bits Bits Bits Bits Bits Bits Bits Bits Bits
9-10 16-29 30-31 Bits
9-10 16-29 30-31 Bits
9-10 16-29 30-31 Bits
9-10 16-29 30-31 Bits Bits Bits Bits Bits
CS1_CS3_7730_011
LRU/System Operations View
RETURN TO RET LRU/SYS LR RU OPS PS S
Figure 30: Review Vibration Data Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
PW1500G Trim Balance Helper Tool Using the tool instructions provided in the Excel off-line tool attached to the Special Instruction, calculate fan trim balance weight(s) and location information as follows: 1. Use Label 47 Fan Trim Balance Solution Weight_FANWGT and Label 54 Fan Trim Balance Solution Angle_FANANG values to determine the list of part numbers and placement based PW1500G Trim Balance Helper Tool. 2. Enter the previously installed trim balance spring loaded weights into the PW1500G Trim Balance Helper Tool. 3. Enter the Label 47 Fan Trim Balance Solution Weight_FANWGT value from the OMS in the Magnitude field of the PW1500G Trim Balance Helper Tool. 4. Enter the Label 54 Fan Trim Balance Solution Angle_FANANG value from the OMS in the Direction field of the PW1500G Trim Balance Helper Tool. 5. Click SET button. 6. Record results on the recommendation screen. 7. Remove or install fan trim balance weight(s) by size and location specified in the PW1500G Trim Balance Helper Tool. 8. Ensure weights are installed with spring forward. 9. Install the inlet cone cover to the inlet cone. Perform Test 8 - Vibration Analysis as per the Aircraft Maintenance Publication. Observe the engine vibration data during engine run with the OMS Further fan trim balance is not necessary once Nf vibration is under 0.5 IPS or 1.8 units. If the vibration survey data indicates that the fan vibration levels are unacceptable, repeat the procedure.
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TECHNICAL TRAINING MANUAL
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77 - Engine Indicating
CS1_CS3_7730_013
77-32 Vibration Monitoring
Figure 31: PW1500G Trim Balance Helper Tool Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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ATA 78 - Exhaust
BD-500-1A10 BD-500-1A11
78 - Exhaust
Table of Contents 78-36 Thrust Reverser Structure ............................................78-2 General Description ..........................................................78-2 Component Location ........................................................78-6 Inner Fixed Structure ...................................................78-6 Translating Sleeve .......................................................78-6 Component Information ....................................................78-8 Translating Sleeves .....................................................78-8 Blocker Doors ............................................................78-10 Cascades ...................................................................78-12 78-31 Thrust Reverser Actuation System.............................78-14 General Description ........................................................78-14 Component Location ......................................................78-16 Isolation Control Unit..................................................78-16 Directional Control Unit ..............................................78-16 Locking Feedback Actuators......................................78-18 Locking Actuators ......................................................78-18 Flexible Drive Shaft/Deploy Tubes.............................78-18 Manual Drive Units.....................................................78-18 Track Lock Valves......................................................78-18 Track Lock Units ........................................................78-18 Component Information ..................................................78-20 Isolation Control Unit..................................................78-20 Directional Control Unit ..............................................78-20 Locking Feedback Actuator ......................................78-22 Locking Actuator .......................................................78-24 Locking Feedback Actuator and Locking Actuator Installation..................................................................78-26 Track Lock Unit ..........................................................78-28 Track Lock Valve .......................................................78-30 Throttle Quadrant Assembly ......................................78-32 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Detailed Component Information ................................... 78-34 Locking Feedback Actuator and Locking Actuator ...................................................... 78-34 Throttle Quadrant Assembly Thrust Lever Control ................................................. 78-36 Controls and Indications ................................................ 78-38 Controls .................................................................... 78-38 Indications ................................................................ 78-40 Detailed Description ...................................................... 78-42 Thrust Reverser Locked............................................ 78-42 Isolation Control Unit Energized Actuators Overstow................................................... 78-44 Track Lock units Released........................................ 78-46 DCU Energized ......................................................... 78-48 Thrust Reversers Fully Deployed .............................. 78-50 Stow Command......................................................... 78-52 Thrust Reverser Stowed ........................................... 78-54 Thrust Reverser Actuating System Control Electrical Schematic ..................................... 78-56 Thrust Reverser Actuating System Control Logic ............................................................. 78-58 Monitoring and Tests ..................................................... 78-60 CAS Messages ......................................................... 78-61 Practical Aspects ............................................................ 78-62 Lockout the Thrust Reversers for Dispatch .............. 78-62 Manually Operate the Translating Sleeves for Maintenance ........................................................ 78-66 Thrust Reverser Cycling Test.................................... 78-68
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_M5_78_ExhaustTOC.fm
78 - Exhaust
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7: Figure 8:
Thrust Reverser Doors .........................................78-3 Thrust Reverser Operation....................................78-5 Thrust Reverser Component Location ..................78-7 Translating Sleeves and Sliders............................78-9 Blocker Door Fittings ...........................................78-11 Cascades ............................................................78-13 Thrust Reverser Actuation System......................78-15 Isolation Control Unit and Directional Control Unit Location.........................78-17 Figure 9: Thrust Reverser Actuation System Component Location ...........................................78-19 Figure 10: Isolation Control Unit and Directional Control Unit .......................................78-21 Figure 11: Locking Feedback Actuator................................78-23 Figure 12: Locking Actuator..................................................78-25 Figure 13: Locking Feedback Actuator and Locking Actuator Installation ...............................78-27 Figure 14: Track Lock Unit ...................................................78-29 Figure 15: Track Lock Valve.................................................78-31 Figure 16: Throttle Quadrant Assembly................................78-33 Figure 17: Locking Feedback Actuator Description..............78-35 Figure 18: Throttle Quadrant Assembly Thrust Lever Control............................................78-37 Figure 19: Thrust Reverser Control ......................................78-39 Figure 20: Thrust Reverser Indications ................................78-41 Figure 21: Thrust Reverser Locked ......................................78-43 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 22: Isolation Control Unit Energized – Actuators Overstow ............................................ 78-45 Figure 23: Track Lock Units Released................................. 78-47 Figure 24: DCU Energized................................................... 78-49 Figure 25: Thrust Reversers Fully Deployed ....................... 78-51 Figure 26: Stow Command .................................................. 78-53 Figure 27: Thrust Reverser Stowed ..................................... 78-55 Figure 28: Thrust Reverser Actuating System Control Electrical Schematic............................... 78-57 Figure 29: Thrust Reverser Actuating System Logic ........... 78-59 Figure 30: Isolation Control Unit Deactivation...................... 78-63 Figure 31: Thrust Reverser Latch Beam Lockout Pins ........ 78-65 Figure 32: Manually Deploy the Thrust Reverser Sleeves .................................... 78-67 Figure 33: Thrust Reverser Cycling Test ............................. 78-69
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-ii
CS130_M5_78_ExhaustLOF.fm
78 - Exhaust
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CS130_M5_78_ExhaustLOF.fm
EXHAUST - CHAPTER BREAKDOWN EXHAUST CHAPTER BREAKDOWN
Thrust Reverser Structure
1 Thrust Reverser Actuation and Control
2
Thrust Reverser Indications
3
78 - Exhaust 78-36 Thrust Reverser Structure
78-36 THRUST REVERSER STRUCTURE GENERAL DESCRIPTION The PW1500G engine has a cascade-type thrust reverser (T/R). Two thrust reverser doors make up the thrust reverser structure. The thrust reverser doors enclose the engine core section and provide the fan bypass airflow path for forward thrust. The composite inner fixed structure (IFS) provides ventilation, fire suppression, and noise attenuation of the engine core. A torque box, mounted on the leading edge of the IFS, provides the mounting points for the actuators, the cascades, the blocker doors and drag links. The outer section of the thrust reverser door houses the translating sleeves. The translating sleeves deploy during landing to direct fan airflow forward and through the cascades to slow the aircraft. Each thrust reverser door is attached to the pylon at the hinge beam. The latch beam has latches that secure the doors closed.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Hinge Beam
Blocker Door
Cascades Torque Box
Inner Fixed Structure Latch Beam
NOTE Thrust reverser shown in deployed position.
CS1_CS3_7833_001
Translating Sleeve
Figure 1: Thrust Reverser Doors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-3
CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Thrust Reverser Operation The inner fixed structure (IFS) and the translating sleeve form the thrust reverser (T/R) door. Each T/R door has five composite blocker doors mounted on the translating sleeve. The upper and lower doors are unique in size and cannot be installed at any other position. The blocker doors lay flush with the translating sleeve when in the stowed position. When the T/R is deployed, the drag links pull the blocker doors into the T/R door. This blocks the fan air from being directed out the exhaust. The airflow is directed in a controlled manner through the cascades in a forward and outward direction. The forward component of the redirected airflow provides the reverse thrust action which aids in slowing the aircraft. The outward component of the airflow is designed to ensure the reverse thrust air is not redirected into the engine inlet, or against the fuselage and flight control surfaces.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Upper Blocker Door
Drag Links
Drag Links
Translating Sleeve Inner Fixed Structure
BLOCKER DOORS IN STOWED POSITION
Lower Blocker Door BLOCKER DOORS IN DEPLOYED POSITION
CS1_CS3_7833_004 S3 7833 004
Blocker Doors
Figure 2: Thrust Reverser Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
COMPONENT LOCATION
TRANSLATING SLEEVE
The thrust reverser structure has the following components:
The translating sleeves are mounted on the IFS at the hinge beam and the latch beam.
•
-
Inner fixed structure (IFS)
-
Hinge beam
Blocker Doors
-
Latch beam
The 10 blocker doors are mounted on the translating sleeves.
-
Torque box
-
Cascades
Translating sleeve -
Blocker doors
INNER FIXED STRUCTURE The inner fixed structure (IFS) encloses the engine core section. Hinge Beam The hinge beams are located on the top of the IFS. Latch Beam The latch beams are located at the bottom of the IFS. Torque Box The torque box is located at the front of the IFS. Cascades The cascades are mounted on the torque box.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_ComponentLocation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Translating Sleeve
Hinge Beam
Cascades
Torque Box Inner Fixed Structure
Latches
Latch Beam
CS1_CS3_7833_002
Blocker Doors
Figure 3: Thrust Reverser Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_78_L2_TRS_ComponentLocation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
COMPONENT INFORMATION TRANSLATING SLEEVES The translating sleeves deploy aft guided by the primary and secondary sliders that move inside the primary and secondary Teflon coated slider tracks, which are attached to the inner fixed structure (IFS). Leading edge bumpers rest on lands on the torque box in order to maintain the aerodynamic contour of the translating sleeve when stowed. The forward section of the inner sleeve have cutouts in which the blocker doors rest flush to provide a streamlined fan bypass air flow.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Primary Slider Track Secondary Slider Track Secondary Slider Blocker Doors Primary Slider
Lands
Leading Edge Bumper
Secondary Slider Track Torque Box
Secondary Slider
INTERIOR VIEW OF TRANSLATING SLEEVE
CS1_CS3_7833_003
Primary Slider
Primary Slider Track
Figure 4: Translating Sleeves and Sliders Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-9
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
BLOCKER DOORS The blocker doors have hinge fittings that mate with clevises attached to the inner wall of the translating sleeve. The doors rotate on these hinge fittings when the sleeves translate rearward. Drag links aid in rotating the blocker doors into the T/R door. The drag links are fixed-length aluminum rods with an airfoil cross-section. Teflon coated spherical bearings on both ends of the drag links ensure smooth movement. An aluminum drag link housing bracket built into the blocker door provides the attachment point for the drag link. A leaf spring mechanism absorbs drag link forces on the blocker door. Rest pads, mounted on supports, contact the blocker doors to absorb vibrations of the doors when stowed. Each blocker door has drain holes to prevent water accumulation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-10
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Blocker Door Seal Clevis
Hinge Fitting Leaf Spring
Drain Holes
Drag Link Housing
Drag Link
Clevis Hinge Fitting
CS1_CS3_7833_005
Spherical Bearings Rest Pad
Figure 5: Blocker Door Fittings Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-11
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
CASCADES The cascades are made up of sixteen individual carbon fiber cascade boxes, designed to turn fan airflow into forward and sideways directions. The cascade boxes are designed to provide specific airflow for their location on the engine. The cascade boxes have fastener patterns that ensure they can not be installed incorrectly. Each cascade box mounts between the torque box and the aft cascade ring. Flow blockers are installed at the hinge beam and latch beam to prevent the airflow from leaking between the cascade boxes and the beams.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-12
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Flow Blocker
1
2
16 3
4
15
5
14
Aft Cascade Ring 6
13
7 12
Torque Box
8
10
9 CASCADE BOX LOCATION
CS1_CS3_7833_006
11
Figure 6: Cascades Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-13
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
78-31 THRUST REVERSER ACTUATION SYSTEM GENERAL DESCRIPTION The reverse actuation system has four hydraulic actuators, two track lock units, an isolation control unit (ICU), and a directional control unit (DCU). The thrust reverser actuation system (TRAS) is controlled through the electronic engine control (EEC), and the control and distribution cabinets (CDCs). The thrust reverser actuation system operates when the aircraft is on the ground. It is controlled by moving the thrust levers on the throttle quadrant assembly (TQA) to the reverse idle position. Each thrust reverser door has a translating sleeve that is operated by a locking feedback actuator and a locking actuator. The operating of the two actuators is synchronized to ensure proper movement of the translating sleeve. The locking feedback actuator has a linear variable differential transformer (LVDT) for position feedback. The actuators have an internal hydraulically released mechanical lock. A proximity switch monitors the lock status of each actuator. Two track lock units lock the translating sleeve when the thrust reverser is stowed. They protect against inadvertent thrust reverser deployment. Each track lock unit is controlled by a track lock solenoid valve. The track lock solenoid valve is controlled by the CDC. The track lock unit releases when the translating sleeve is ready to be deployed. The track lock unit relocks the translating sleeve once it stows. The thrust reverser actuation system (TRAS) has an isolation control unit (ICU) that isolates the system from supply pressure when the system is non-operational and a directional control unit (DCU) that controls the directional functions of the thrust reverser translating sleeves. The ICU and DCU have solenoid valves that are controlled by the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
To allow for dispatch of inoperative thrust reversers (T/Rs), and to provide the additional safety during maintenance activities, a manual inhibit lever is installed on the ICU to isolate the thrust reverser actuation system (TRAS). The DCU has a directional spool valve operated by a directional control solenoid valve that controls the unlock, deploy, stow and lock sequence when commanded. The thrust reverser (T/R) is deployed by moving the thrust levers from the forward thrust position to the idle position and then to the reverse idle position. Once the translating sleeves have reached 85% of full travel, the thrust levers can be moved to the maximum reverse thrust position. The EEC provides a higher maximum reverse thrust for a rejected takeoff than for a rollout after landing. The T/Rs are stowed by returning the thrust levers to the idle position. The thrust levers can not be moved beyond the idle position and engine thrust is limited until the thrust reversers are fully stowed. During the deploy sequence, a white REV icon is displayed on EICAS. When the thrust reverser (T/R) is fully deployed, the REV icon turns green. When the T/Rs are stowed, the white REV icon is displayed until the T/R is fully locked. The locking feedback actuator, locking actuator, and track lock unit provide mechanical protection against an inadvertent thrust reverser (T/R) deployment. The EEC controlled ICU and DCU, and the CDC controlled track lock solenoid valves provide electrical protection against an inadvertent thrust reverser (T/R) deployment. The EEC monitors the T/R position through the locking feedback actuator LVDTs, lock status, and the track lock unit lock status. If a thrust reverser inadvertently deploys, the EEC reduces the engine thrust to idle.
TECHNICAL TRAINING MANUAL
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78-14
CS130_78_L2_TRAS_General.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND
ISOLATION CONTROL UNIT
REV
Air/Gnd
TRACK LOCK VALVE
ELECTRONIC ENGINE CONTROL Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
CDC
DMC 1 and DMC 2
SOL
CDC DMC LVDT SOL
Pressure Return Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor ARINC 429
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_032
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 7: Thrust Reverser Actuation System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRAS_General.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
COMPONENT LOCATION The thrust reverser actuating system has the following components: •
Isolation control unit (ICU)
•
Directional control unit (DCU)
•
Locking feedback actuators (Refer to Figure 9)
•
Locking actuators (Refer to Figure 9)
•
Flexible drive shaft/deploy tube (Refer to Figure 9)
•
Manual drive units (Refer to Figure 9)
•
Track lock valves (Refer to Figure 9)
•
Track lock units (Refer to Figure 9)
ISOLATION CONTROL UNIT The isolation control unit (ICU) is located in the aft pylon section and is accessible from access panels on each side of the pylon. DIRECTIONAL CONTROL UNIT The directional control unit (DCU) is located in the forward pylon section, above the engine core section area.
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CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
A
B
B DIRECTIONAL CONTROL UNIT
CS1_CS3_7832_005
A ISOLATION CONTROL UNIT
Figure 8: Isolation Control Unit and Directional Control Unit Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-17
CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING FEEDBACK ACTUATORS The locking feedback actuators are located on each thrust reverser door upper section. LOCKING ACTUATORS The locking actuators are located on each thrust reverser door lower section. FLEXIBLE DRIVE SHAFT/DEPLOY TUBES The flexible drive shaft/deploy tubes are located on the torque box between the locking actuator, and the locking feedback actuator. MANUAL DRIVE UNITS The manual drive units are located on each of the locking actuators. TRACK LOCK VALVES The track lock valves are located on each of the thrust reverser door lower section, near the latch beam. TRACK LOCK UNITS The track lock units are located on each of the thrust reverser latch beam.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-18
CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Locking Feedback Actuator
Return Deploy Stow Flexible Drive Shaft/Deploy Tubes
Locking Actuator
Track Lock Valves
Manual Drive Units
Locking Actuator
CS1_CS3_7832_002
Hydraulic Line
Track Lock Units
Figure 9: Thrust Reverser Actuation System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-19
CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
COMPONENT INFORMATION ISOLATION CONTROL UNIT The isolation control unit (ICU) isolates the thrust reverser actuating system (TRAS) from the aircraft hydraulic system. The ICU has a dual-channel solenoid controlled by either electronic engine control (EEC) channel. The solenoid operates a pilot valve that supplies hydraulic pressure to an internal isolation spool valve. The isolation spool valve opens to allow hydraulic pressure from the aircraft hydraulic system to enter the TRAS. To allow for dispatch of an inoperative thrust reverser (T/R) and to provide additional safety during ground maintenance activities, a manual inhibit lever is installed to prevent the isolation spool valve from operating. An inhibit pin stored on the ICU locks the manual inhibit lever in the inhibit position. The position of the inhibit lever is monitored by dual-channel inhibit proximity sensors. DIRECTIONAL CONTROL UNIT The directional control unit (DCU) has a dual-channel solenoid controlled by either EEC channel. The solenoid operates a directional control valve that supplies hydraulic pressure to an internal directional spool valve. The directional spool valve provides hydraulic pressure to the actuators and the track lock valves to control unlocking, deploying, stowing, and locking of the thrust reversers (T/Rs). A dual-channel proximity pressure sensor provides the status of the hydraulic pressure at the outlet of the ICU to each channel of the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Dual-Channel Solenoid
Dual-Channel Inhibit Proximity Sensors Inhibit Pin
ISOLATION CONTROL UNIT
Manual Inhibit Lever
Dual-Channel Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT
CS1_CS3_7832_034
Dual-Channel Solenoid
Figure 10: Isolation Control Unit and Directional Control Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-21
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING FEEDBACK ACTUATOR The locking feedback actuator consists of a jackscrew that drives the translating sleeve. The jackscrew is hydraulically operated. The stow tube supplies hydraulic fluid to stow the translating sleeve. The rod end fitting at the aft end of the actuator provides a connection point to the translating sleeve. The actuator has an internal lock that maintains the actuator in the stowed position when hydraulic pressure is not applied. The lock releases when hydraulic pressure is applied. This allows the translating sleeve to deploy. When the translating sleeve stows, the lock re-engages. A proximity sensor monitors the lock and provides a stowed signal to the electronic engine control (EEC). The actuator has a manual lock that mechanically unlocks the actuator to allow the translating sleeve to be manually deployed. The locking feedback actuator connects to the locking actuator through a flexible drive shaft and deploy tube. The flexible drive shaft synchronizes the movement of the two actuators. The deploy tube supplies hydraulic fluid to the locking actuator. A linear variable differential transformer (LVDT), mounted on the forward end of the actuator, provides position feedback of the translating sleeves to the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-22
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Rod End
Stow Tube
Manual Lock
Jackscrew Shaft
Proximity Sensor Dual-Channel Linear Variable Differential Transformer
Flexible Drive Shaft
CS1_CS3_7832_033
Deploy Tube
Figure 11: Locking Feedback Actuator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-23
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING ACTUATOR The locking actuator is similar in construction to the locking feedback actuator except it does not have an LVDT and its position is not monitored. The actuator has an internal lock that maintains the actuator in the stowed position when hydraulic pressure is not applied. The lock releases when hydraulic pressure is applied. This allows the translating sleeve to deploy. When the translating sleeve stows, the lock re-engages. A proximity sensor monitors the lock and provides a stowed signal to the EEC. The actuator has a manual lock that mechanically unlocks the actuator to allow the translating sleeve to be manually deployed. The locking actuator connects to the locking feedback actuator through a flexible drive shaft and deploy tube. The flexible drive shaft synchronizes the movement of the two actuators. The deploy tube carries hydraulic fluid to the actuator. A manual drive unit (MDU) mates with the actuator flexible drive shaft. This rotational input drives the locking actuator through the flex drive shaft, and the locking feedback actuator. A torque limiter in the MDU prevents over torquing of the actuator during manual operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Flexible Drive Shaft
Deploy Tube
CS1_CS3_7832_009 32_009
Manual Lock
Proximity Sensor Manual Drive Unit
Figure 12: Locking Actuator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-25
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING FEEDBACK ACTUATOR AND LOCKING ACTUATOR INSTALLATION The locking feedback actuator and locking actuators are mounted in gimbals on the torque box. The gimbal allow the actuator to pivot as the translating sleeve deploys. This prevents the translating sleeve from jamming. The rod end fitting at the aft end of the actuator is attached to the translating sleeve. A panel provides access to the rod end fitting.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-26
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Rod End Fitting
A
Translating Sleeve
B
A ROD END OF ACTUATOR SHOWN WITH ACCESS PANEL REMOVED
Gimbal
B
HYDRAULIC LOCKING FEEDBACK ACTUATOR
CS1_CS3_7832_006
Torque Box
Figure 13: Locking Feedback Actuator and Locking Actuator Installation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-27
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK UNIT The track lock unit is installed on the latch beam of each thrust reverser door. The track lock unit is operated by hydraulic pressure supplied through the track lock valves. The track lock unit has a spring-loaded bolt that protrudes into the path of the translating sleeve. In the event of inadvertent deployment, the track lock unit prevents the translating sleeve from deploying. The bolt retracts when hydraulic pressure is supplied. Two proximity sensors monitor the position of the track lock unit bolt and provide the lock status to the EEC. An unlock lever is used to retract the bolt when the translating sleeve is manually deployed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-28
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Track Lock Bolt
Track Lock Unit
Proximity Sensors
Translating Sleeve
CS1_CS3_7832_012
Unlocking Lever
Figure 14: Track Lock Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-29
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK VALVE The track lock valve is a solenoid powered valve used to supply hydraulic pressure to unlock the track lock unit. When the translating sleeve deploys, the track lock valve solenoid stays energized. This allows the track lock unit to remain unlocked until the thrust reverser is returned to the stowed position.
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78-30
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CS1_CS3_7832_015
Solenoid Valve
Figure 15: Track Lock Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-31
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THROTTLE QUADRANT ASSEMBLY The throttle quadrant assembly (TQA) provides thrust lever position to the EEC. The EEC calculates the engine power required between REV and the MAX REV position 19.5° aft of IDLE. Each thrust lever has two thrust reverser stow/deploy switches that provide thrust lever positions to the control and distribution cabinet (CDC). These switches are activated by cams at 1.85° aft of thrust lever idle position. The TQA baulk mechanism prevents the thrust levers from being moved beyond 7.5° aft of IDLE until the thrust reverser translating sleeves are fully deployed. The baulk mechanism prevents the thrust levers from moving 2.5° forward of idle position until the translating sleeves are fully stowed.
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78-32
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Forward Thrust Baulk Release
IDLE
Reverse Switch Activation REV Idle Detent
1.85°
4.6°
2.5°
Reverse Thrust Baulk Release
Baulk Mechanism
MAX REV 7.5° 19.5°
Thrust Reverser Switches
Cams
CS1_CS3_7600_016
THROTTLE QUADRANT ASSEMBLY
Figure 16: Throttle Quadrant Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-33
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
DETAILED COMPONENT INFORMATION LOCKING FEEDBACK ACTUATOR AND LOCKING ACTUATOR When hydraulic pressure is supplied from the isolation control unit (ICU), it is routed through the stow tube, to the stow port of the actuator, keeping the actuator stowed and locked. When the thrust reverse is selected, hydraulic pressure, supplied through the deploy port is directed to the deploy side of the actuator. The hydraulic pressure pushes the lock sleeve clear of the lock tine, allowing the jackscrew to rotate. The hydraulic pressure acts on the internal side of the jackscrew. Both sides of the actuator are now pressurized. Since the jackscrew surface area is greater than the stow side of the actuator piston, the actuator deploys. When the thrust reversers (T/Rs) are commanded to stow, the deploy side of the actuator is directed to the return port and the hydraulic pressure on the stow side drives the actuator to the stow position. While deploying or stowing, the jackscrew worm wheel turns the flexible drive shaft worm gear. The flexible drive shaft worm gear is connected to the flexible drive shaft that interconnects the locking feedback actuator to the locking actuator. This allows the actuators to maintain synchronization during the stow or deploy cycle. The manual lock is used to move the lock sleeve when the translating sleeve is manually deployed. The LVDT on the locking feedback actuator is integral to the actuator and is not a line replaceable unit (LRU). The LVDT has an internal self-rigging mechanism that operates on the first deploy and stow cycle. The self-rigging mechanism establishes the stowed and locked position for the actuator. All position feedback signals to the engine electronic control (EEC) are referenced from the actuator stowed and locked position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-34
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Stow Tube
Manual Lock
Jack Screw Shaft
Dual-Channel Linear Variable Differential Transformer
Flexible Drive Shaftr
A
Gimbal Manual Lock Lock Sleeve
Actuator Piston
Worm Wheel
Deploy Port
Jackscrew
B Flexible Drive Shaft Input
Stow Hydraulic Pressure
A
Deploy Hydraulic Pressure
Lock Return Tine Port
Worm Wheel
Stow Port
CS1_CS3_7832_007
B
Figure 17: Locking Feedback Actuator Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-35
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THROTTLE QUADRANT ASSEMBLY THRUST LEVER CONTROL The gate plate, mounted on the TQA, has end stops at the IDLE, MAX and MAX REV positions. A gate pin on the thrust reverse link contacts the end stops to limit the movement of the thrust lever. An IDLE stop prevents the thrust lever from moving into the thrust reverse range. Thrust reverser finger lifts on each thrust lever raise the gate pin clear of the IDLE stop. When the gate pin is clear of the IDLE stop, the thrust lever can be moved into the thrust reverse range. The thrust lever can be returned to the forward thrust range without using the finger lifts. When the baulk solenoid energizes, it positions the baulk mechanism rocker arms to lock the thrust lever in the idle range until the thrust reverser is either fully deployed or fully stowed. If required, the interlock can be overridden with a force of approximately 25 lb.
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78-36
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Baulk Mechanism
Thrust Lever
B Thrust Reverser Finger Lifts
Rocker Arms
A
Thrust Reverser Link
IDLE Stop MAX Stop
Gate Pin
Gate Plate
A THRUST REVERSER BAULK MECHANISM (PLATE REMOVED FOR CLARITY)
B THRUST LEVER (COVER REMOVED FOR CLARITY)
CS1_CS3_7600_005 S1 CS3 7600 005
Baulk Solenoid
MAX REV Stop
Figure 18: Throttle Quadrant Assembly Thrust Lever Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-37
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CONTROLS AND INDICATIONS CONTROLS The thrust reversers (T/Rs) are controlled from the throttle quadrant assembly (TQA). The thrust levers are moved into the reverse thrust range using the finger lifts located on the thrust levers. The thrust lever is used to control the amount of thrust required when the T/R has deployed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-38
CS130_78_L2_TRAS_ControlsIndications.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CS1_CS3_7831_001
Thrust Reverser Finger Lifts
Figure 19: Thrust Reverser Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-39
CS130_78_L2_TRAS_ControlsIndications.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
INDICATIONS The thrust reversers position is displayed on the engine indication and crew alerting system (EICAS) N1 gauge. The REV icon shows the status of the thrust reverser. •
A green REV icon appears in the N1 analog gauge when the thrust reversers (T/Rs) are fully deployed
•
A white icon indicates the thrust reversers are in transition
•
A amber REV icon indicates a T/R fault when the aircraft is on the ground, the thrust reverser is not manually inhibited, and any one of the following conditions occurs: -
Inadvertent thrust reverser deployment
-
Thrust reverser translating sleeve position feedback failed on either translating sleeve
-
Thrust reverser is jammed
-
Thrust reverser inability to deploy
-
Thrust reverser inadvertently pressurized
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CS130_78_L2_TRAS_ControlsIndications.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
REV
N1
70.4 REV TABLE 1 EICAS FLAG Symbol
688
692
REV
Aircraft is on ground, the thrust reverser is not manually inhibited, and any one of the following conditions: Ɣ7KUXVWUHYHUVHULQDGYHUWHQWGHSOR\PHQW Ɣ7KUXVWUHYHUVHUWUDQVODWLQJVOHHYHSRVLWLRQIHHGEDFN failed on either translating sleeve Ɣ7KUXVWUHYHUVHULVMDPPHG Ɣ7KUXVWUHYHUVHULQDELOLW\WRGHSOR\ Ɣ7KUXVWUHYHUVHULQDGYHUWHQWO\SUHVVXUL]HG
REV
On ground: thrust reverser in transition.
REV
On ground: thrust reverser is deployed.
EGT
87.5 1490 105 110
N2 FF (PPH) OIL TEMP OIL PRESS
87.5 1495 104 110
Description
NOTE The thrust reverser icon has priority over the vibration icon on the ground.
CS1_CS3_7837_005
70.5
Figure 20: Thrust Reverser Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-41
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78 - Exhaust 78-31 Thrust Reverser Actuation System
DETAILED DESCRIPTION THRUST REVERSER LOCKED When the thrust reverser (T/R) is stowed and the thrust levers are in forward thrust, the isolation control unit (ICU) and directional control unit (DCU) isolate hydraulic pressure from the thrust reverser system. The ICU isolation control valve control solenoid is de-energized. The directional control unit (DCU) pressure proximity sensors monitor the hydraulic pressure in the thrust reverser actuating system. The track lock valve solenoids are also de-energized and the track lock units are in the locked position. The track lock proximity sensors monitor the track lock units status. With the DCU directional control valve solenoid de-energized, the actuator is locked in the stowed position. The locking actuator proximity sensors monitor the lock mechanism position to ensure the thrust reverser is stowed and locked. The feedback actuator LVDTs provide position feedback as a secondary indication of the actuator position.
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78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND
RH Track Lock Position LH Track Lock Position
SOL
LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Linear Variable Differential Transformer Solenoid Proximity Sensor
TRACK LOCK VALVE
ICV Command ICU Pressure
DIRECTIONAL CONTROL UNIT
SOL
ISOLATION CONTROL UNIT Isolation Control Valve
Pressure Proximity Sensor
SOL
Directional Control Valve
Track Lock Unit
LOCKING ACTUATOR LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
NOTE Left engine reverser shown.
Track Lock Unit CS1_CS3_7832_018
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 21: Thrust Reverser Locked Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-43
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ISOLATION CONTROL UNIT ENERGIZED - ACTUATORS OVERSTOW When reverse thrust is selected, the electronic engine control (EEC) energizes the ICU isolation valve control solenoid. The hydraulic pressure is directed through the isolation spool valve to the DCU. The dual pressure proximity sensor, installed on the DCU, signals the EEC that hydraulic pressure is available from the ICU. Hydraulic pressure is supplied to the stow side of the locking feedback actuators and the locking actuators. The actuators over stow to release the locking mechanisms. The proximity sensors on the actuators indicate the locking mechanism is in the locked position. Hydraulic pressure is also supplied to the track lock valves. The track lock units remain locked. The track lock unit proximity sensors send the locked status to the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-44
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND SOL
LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Linear Variable Differential Transformer Solenoid Proximity Sensor
RH Track Lock Position
TRACK LOCK VALVE
LH Track Lock Position ICV Command ICU Pressure
ISOLATION CONTROL UNIT
DIRECTIONAL CONTROL UNIT
SOL
SOL
Isolation Control Valve
Pressure Proximity Sensor
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
NOTE Left engine reverser shown.
Track Lock Unit CS1_CS3_7832_019
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 22: Isolation Control Unit Energized – Actuators Overstow Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-45
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK UNITS RELEASED With the WOW logic satisfied, the control and distribution cabinet (CDC) energizes the track lock valve solenoids open both track lock valves. The hydraulic pressure is now supplied to the track lock unit and the track lock unit bolt retracts. The track lock unit proximity sensors signal the EEC indicating that the track lock unit bolts are clear of the translating sleeve and the thrust reverser is ready to deploy.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-46
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CDC TRACK LOCK VALVE COMMAND
LEGEND
RH Track Lock Position LH Track Lock Position
TRACK LOCK VALVE
SOL
CDC LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
ICV Command ICU Pressure
ISOLATION CONTROL UNIT
DIRECTIONAL CONTROL UNIT
SOL
SOL
Isolation Control Valve
Pressure Proximity Sensor
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_020
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 23: Track Lock Units Released Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-47
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
DCU ENERGIZED The EEC energizes the DCV directional control valve solenoid. This positions the directional spool valve to pressurize the actuators deploy side to completely unlock the actuators. The actuator proximity sensors and LVDTs provide the unlock status to the EEC and the translating sleeves start to deploy. The track lock valves solenoids remain energized during the deploy sequence. A white REV icon is displayed on the EICAS N1 gauge.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-48
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ELECTRONIC ENGINE CONTROL
LEGEND
ISOLATION CONTROL UNIT
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position LH Sleeve 95% Deployed RH Sleeve 95% Deployed
TRACK LOCK VALVE
SOL
CDC LVDT SOL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
CDC TRACK LOCK VALVE COMMAND
REV
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_021
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 24: DCU Energized Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-49
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSERS FULLY DEPLOYED The translating sleeve position is sent to the EEC by the locking feedback actuator LVDTs. When the translating sleeves reach 85% of full travel, the thrust levers may be moved to the full reverse thrust. When the translating sleeves reach 95% of full travel, the green REV icon is shown on the N1 gauge.
NOTE Both sides of the actuators are pressurized, however the greater actuator piston area on the deploy side ensures the actuators deploy.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-50
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ELECTRONIC ENGINE CONTROL
LEGEND
ISOLATION CONTROL UNIT
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position LH Sleeve 95% Deployed RH Sleeve 95% Deployed
TRACK LOCK VALVE
SOL
CDC LVDT SOL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
CDC TRACK LOCK VALVE COMMAND
REV
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_022
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 25: Thrust Reversers Fully Deployed Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-51
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
STOW COMMAND When the throttle moves from reverse idle to forward idle, the EEC de-energizes the DCU control solenoid. The directional spool valve repositions and the deploy side of the actuators are connected to the return side of the hydraulic system. Since hydraulic pressure always remain on the stow side of the actuators, the translating sleeves stow. The locking feedback actuator LVDTs feedback to the EEC changes the REV icon from green to white.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-52
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ELECTRONIC ENGINE CONTROL
LEGEND
ISOLATION CONTROL UNIT
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position LH Sleeve 95% Deployed RH Sleeve 95% Deployed
TRACK LOCK VALVE
SOL
CDC LVDT SOL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
CDC TRACK LOCK VALVE COMMAND
REV
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_023
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 26: Stow Command Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-53
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER STOWED The track lock valve solenoids de-energize, and the track lock units relock. When the actuator proximity sensors indicate the actuators are in the stowed position, the EEC de-energize the track lock valve solenoids. The spring-loaded track lock unit bolts extend, locking the translating sleeves. The track lock unit proximity sensors indicate the track lock unit lock status to the EEC. With the confirmation of the track lock units in the locked position, the EEC de-energizes the ICU control solenoid, isolating the trust reverser actuating system from the hydraulic system. The thrust reverser (T/R) hydraulic components are connected to the return line. The DCU pressure proximity sensors confirm the pressure is not available to the thrust reverser actuating system (TRAS). The REV indication is removed from the N1 gauge.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-54
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT
SOL
ISOLATION CONTROL UNIT Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position
SOL
Directional Control Valve
SOL
LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Linear Variable Differential Transformer Solenoid Proximity Sensor
TRACK LOCK VALVE
Track Lock Unit
LOCKING ACTUATOR LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_024
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 27: Thrust Reverser Stowed Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-55
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER ACTUATING SYSTEM CONTROL ELECTRICAL SCHEMATIC
The track lock units have a proximity sensor to send a signal to the EEC when the track lock units are unlocked.
The thrust levers in the throttle quadrant assembly (TQA) provide the command signal to the EEC to control the thrust reverser actuating system (TRAS). When the thrust lever is moved to the REV IDLE position, the stow/deploy switches are made. The stow/deploy switches send a discrete to the thrust reverser lock solid-state power controller (SSPC) in control and distribution cabinet 1 (CDC 1).
The EEC controls the TQA baulk solenoid based on LVDTs and the stow and lock proximity sensors feedback received from the actuators. The EEC also monitors the DCU status. The LVDTs are dual-channel with one output wired to each channel of the EEC.
The baulk solenoid that prevents the thrust lever from moving towards the MAX REV range is powered by the TQA L Reverser Lock SSPC in CDC 3. Once the thrust reversers are deployed beyond 85%, the EEC provides a ground to energize the baulk solenoid. The EEC controls the ICU and DCU control solenoids. The ICU solenoid receives power from DC BUS 1 from CDC 1 on channel A and DC BUS 2 from CDC 2 on channel B. The combinational logic in the CDCs enables them when the stow/deploy switches are made, and one of the following conditions is true: •
Weight-on-wheels (WOW) signal from the landing gear and steering control unit (LGSCU)
•
Wheel spin signal from the brake data concentrator unit (BDCU)
The WOW signal from the LGSCU is hardwired into the CDCs, and sent through the data concentrator unit module cabinets (DMCs) to the EEC. The EEC provides the ground to energize the ICV solenoid. The DCU solenoid receives both power and a ground from the EEC. A dual channel pressure sensor provides the EEC with a signal when hydraulic pressure is flowing through the ICU and into the DCU. Once the ICU and DCU control solenoids become energized, the control valves are actuated which allows hydraulic pressure to flow to the actuators and track lock valves. The track lock valves are controlled by solenoids. These solenoids are energized by DC BUS 1 from CDC 1 when the stow/deploy switches are made and the aircraft is on the ground.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Once the T/Rs are deployed beyond 85%, the EEC energizes the baulk solenoid. This allows the thrust lever to move into MAX REV range. When the thrust reversers deploy beyond 95%, the EEC sends a signal to the EICAS for indicating the green REV icon. Maximum reverse thrust is determined by the EEC. The EEC makes this determination using thrust lever angle (TLA). If the TLA was above 23° and then retarded to idle, the EEC determines that a rejected takeoff has occurred and sets the maximum reverse thrust level at 87.4% N1. During rollout after landing, the maximum reverse thrust is 80% N1. If the airspeed is below 55 kt and reverse thrust is still applied, the maximum reverse thrust N1 is reduced to reverse idle N1. If either thrust lever is moved in the reverse thrust range while the aircraft is in flight, a THROTTLE IN REVERSE caution message is displayed. Inadvertent operation of the thrust reverser (T/R) is detected by the EEC through monitoring the T/R position feedback and the lock status. If the T/R inadvertently deploys, the EEC commands the engine power to idle. An L(R) REVERSER FAIL caution message is displayed if the EEC detects faults that prevent system operation. On the ground, an amber REV icon is also displayed in the N1 indicator. If both locks on a translating sleeve indicate unlocked, a track lock indicates unlocked, or the TRAS is inadvertently pressurized, then a L(R) REVERSER UNLOCK caution message is displayed. The T/R may still be usable for landing. A dual-channel proximity sensor monitors the manual inhibit lever on the ICU. If the thrust reverser (T/R) is locked out, all other T/R CAS messages are inhibited and an L(R) REVERSER INHIBIT status message is displayed.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-56
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK UNIT
CDC 1-10-9 L Thrust Reverser Lock DC SSPC 5A BUS 1
DMC 1 and DMC 2
LGSCU
Logic: Advanced
TQA
CDC 3-4-15 TQA L Reverser Lock DC SSPC 3A BUS 1
LVDT
ACTUATOR Baulk Solenoid ACTUATOR
CDC 1-11-14 L Thrust Reverser ICV A DC SSPC 3A BUS 1
ACTUATOR
EEC
LVDT TRACK LOCK UNIT
Channel A
Logic: Combinational
TRACK LOCK SOLENOID VALVE
MANUAL INHIBIT
Proximity Switch
Channel A Channel B
Channel B
ICU
DCU
ICV ENABLE SOL
DCV ENABLE SOL
Channel A
Channel B
Channel B PRESSURE SENSOR
Return Isolation Control Valve
Hydraulic System 1
LEGEND
Channel A
Channel A
Channel B NOTE Left engine shown. Right engine similar.
Directional Control Valve
BDCU CDC DCU DCV DMC EEC ICU ICV LGSCU LVDT TQA
Pressure Return Brake Data Concentrator Unit Control and Distribution Cabinet Directional Control Unit Directional Control Valve Data Concentrator Unit Module Cabinet Electronic Engine Control Isolation Control Unit Isolation Control Valve Landing Gear and Steering Control Unit Liner Variable Differential Transformer Throttle Quadrant Assembly ARINC 429
CS1_CS3_7832_025
Logic: Combinational
ACTUATOR
Stow Deploy Switches
Logic: Always On
CDC 2-9-11 L Thrust Reverser ICV B DC SSPC 3A BUS 2
Proximity Switch
TRACK LOCK SOLENOID VALVE
BDCU
Figure 28: Thrust Reverser Actuating System Control Electrical Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-57
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER ACTUATING SYSTEM CONTROL LOGIC Isolation Control Valve Logic
Directional Control Valve Logic
The ICV logic requires an aircraft on the ground signal supplied by landing gear and steering control unit (LGSCU) 1 or LGSCU2 WOW signal, or an aircraft touchdown signal supplied by any 2 of the 4 wheel speed units (WSU) indicating a wheel speed of 45 kt.
The DCU logic requires an aircraft on the ground signal supplied by landing gear and steering control unit (LGSCU) 1 or LGSCU2 WOW signal or an aircraft touchdown signal supplied by any left WSU and any right WSU indicating a wheel speed of 45 kt.
The ICV energizes when:
The DCV energizes when:
•
Thrust reverse is selected
•
Thrust reverse is selected
•
No manual inhibit
•
No manual inhibit
•
No thrust reverser fault detected
•
No thrust reverser fault detected
•
Engine running
•
Track lock units unlocked
•
Engine running
The ICV also energizes on the ground when the engine is not running and the EEC thrust reverser cycling test is carried out through the onboard maintenance system (OMS). Track Lock Solenoid Control Logic Once the aircraft is on the ground, the landing gear and steering control unit (LGSCU) 1 or LGSCU 2 supply a WOW signal to the track lock logic. When the thrust levers are selected to the reverse thrust range beyond the -1.85° and both left stow/deploy switches 1 and 2 are actuated by the thrust lever, the track lock solenoid is energized When the thrust reverser is stowed, the track lock solenoids are energized for 11 seconds to provide enough time for the translating sleeve to stow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The DCV also energizes on the ground when the engine is not running and the EEC thrust reverser cycling test is carried out through the onboard maintenance system (OMS). Baulk Solenoid Logic When thrust reverse is selected, each thrust lever is locked in the reverse idle position by the baulk mechanism. The thrust lever is locked until the thrust reverser translating sleeves have deployed far enough to safely increase thrust. When the directional control valve (DCV) has energized and the translating sleeves have deployed beyond 85%, the baulk solenoid energizes and releases the baulk mechanism. The thrust lever can be moved beyond the reverse idle position.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-58
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LGSCU WOW 1 LGSCU WOW 2 2 Out of 4 Wheel Speed Units > 45 Knots
Manual Inhibit No Fault Detected Thrust Reverse Selected
Isolation Control Unit Solenoid Energized
Engine Running Engine Off and EEC Thrust Reverser Cycling Test ISOLATION CONTROL UNIT SOLENOID LOGIC LGSCU WOW 1 LGSCU WOW 2
Track Lock Solenoid Energized Left Stow/Deploy Switch 1 Left Stow/Deploy Switch 2 TRACK LOCK SOLENOID LOGIC
1 of 2 Left Wheel Speed Units > 45 Knots 1 of 2 Right Wheel Speed Units > 45 Knots
LGSCU WOW 1 LGSCU WOW 2
Manual Inhibit No Fault Detected Thrust Reverse Selected Track Locks Unlocked
Directional Control Unit Solenoid Energized
Engine Running Engine Off and EEC Thrust Reverser Cycling Test DIRECTIONAL CONTROL UNIT SOLENOID LOGIC
Baulk Solenoid Energized
Left Translating Sleeve 85% Deployed LEGEND EEC LGSCU WOW
Electronic Engine Control Landing Gear and Steering Control Unit Weight-On-Wheels
Right Translating Sleeve 85% Deployed NOTE
BAULK SOLENOID LOGIC
Left engine shown. Right engine similar.
CS1_CS3_7832_026
Directional Control Unit Solenoid Energized
Figure 29: Thrust Reverser Actuating System Logic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-59
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the thrust reverser system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-60
CS130_78_L2_TRAS_MonitoringTests.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L REVERSER FAIL
Left thrust reverser failed or not available. Icon is only shown if T/R has been commanded.
R REVERSER FAIL
Right thrust reverser failed or not available. Icon is only shown if T/R has been commanded.
L REVERSER UNLOCK
Multiple left reverser locks unlocked.
R REVERSER UNLOCK
Multiple right reverser locks unlocked.
THROTTLE IN REVERSE
Left or right thrust reverser selected in flight.
Table 2: ADVISORY Messages MESSAGE
LOGIC
L ENGINE FAULT Loss of redundant non-critical function for the left engine. R ENGINE FAULT Loss of redundant non-critical function for the right engine.
Table 3: STATUS Messages MESSAGE
LOGIC
L REVERSER INHIBIT
Left T/R manually inhibited as per maintenance procedure (crew awareness).
R REVERSER INHIBIT
Right T/R manually inhibited as per maintenance procedure (crew awareness).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-61
CS130_78_L2_TRAS_MonitoringTests.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
PRACTICAL ASPECTS LOCKOUT THE THRUST REVERSERS FOR DISPATCH The T/Rs can be locked out for dispatch per the Aircraft Minimum Equipment List (MEL). The following is a synopsis of the Aircraft Maintenance Publication (AMP) procedure. Always consult the AMP before beginning any maintenance on the aircraft. There are two steps to deactivate the thrust reversers: •
Deactivate the isolation control unit (ICU)
•
Install a T/R lockout pin on the latch beam assembly
ICU DEACTIVATION The ICU is deactivated by using an inhibit lever located on the ICU, and installing a safety pin to maintain the lever in the inhibited position. The safety pin is stowed on the ICU when not required. Access to the ICU is through access panels on each side of the pylon.
WARNING THE THRUST REVERSER ACTUATION SYSTEM SHOULD BE DEACTIVATED USING THE ICU INHIBIT LEVER, BEFORE PERFORMING ANY MAINTENANCE. THE ACCIDENTAL OPERATION OF THE THRUST REVERSER CAN CAUSE INJURIES TO PERSONNEL, AND DAMAGE TO EQUIPMENT.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-62
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER OPERATIONAL
THRUST REVERSER LOCKED OUT
CS1_CS3_7836_001
REMOVE BEFORE FLIGHT
Lockout Pin Stowage
Figure 30: Isolation Control Unit Deactivation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-63
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Thrust Reverser Latch Beam Lockout Pin After locking out the ICU, lockout pins are installed on the latch beam assemblies to prevent the T/Rs from deploying. The lockout pins are stored on the thrust reverser door latch access panel. The pins are removed from their storage location and installed on the latch beam. The blanking plates that cover the lockout pin holes are stored on the latch access panel when the lockout pins are installed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-64
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
B
A Lockout Pin Blanking Plate
A
Lockout Pins
Latch Access Panel
NOTE Left side shown, right side similar.
B
CS1_CS3_7832_029
Translating Sleeve
Figure 31: Thrust Reverser Latch Beam Lockout Pins Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-65
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Manually Stow the Translating Sleeve
MANUALLY OPERATE THE TRANSLATING SLEEVES FOR MAINTENANCE The locking feedback actuators, locking actuators, and the track lock unit must be unlocked before manually operating the translating sleeves. Manually Deploy the Translating Sleeve Track Lock Unit Unlocking To unlock the track lock unit: •
Move the manual lock from the lock position to the unlock position
•
Install the unlock pin in the manual lockout hole.
Locking Feedback Actuator and Locking Actuator Unlocking To unlock the locking feedback actuator and locking actuator: •
Move the manual lock on the locking feedback actuator to the unlock position and install the unlock pin
•
Move the manual handle on the locking actuator to the unlock position and install the unlock pin
Translating Sleeve Deployment
Stow the Translating Sleeve Manually stow the translating sleeve as follows: •
Put the square drive tool in the manual drive unit
•
Turn the manual drive unit to stow the translating sleeve
•
Put the cover on the manual drive unit
Track Lock Unit Locking To lock the track lock unit: •
Remove the unlock pin from the manual lockout hole
•
Move the manual lock from the unlock position to the lock position
Locking Feedback Actuator and Locking Actuator Locking To lock the locking feedback actuator and locking actuator: •
Remove the unlock pin on the locking feedback actuator and move the manual lock back to the lock position
•
Remove the unlock pin on the locking actuator and move the manual lock back to the lock position
Manually deploy the translating sleeve as follows: •
Remove the cover and put the square drive tool in the manual drive unit
•
Turn the manual drive unit to extend the translating sleeve to the position necessary to do the maintenance
NOTE A maximum speed of 120 rpm is permitted for the manual drive unit. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-66
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Manual Drive Unit
Square Drive Tool Engagement
Locking Actuator
Lock Pin Manual Lock
Manual Lock
Proximity Sensor
Lock Pin
ACTUATOR MANUAL UNLOCKING FEATURE Manual Lockout Hole NOTE TRACK LOCK UNIT IN LOCKED POSITION
Locking actuator shown, locking feedback actuator similar.
CS1_CS3_7832_030
Cover
Figure 32: Manually Deploy the Thrust Reverser Sleeves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-67
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER CYCLING TEST The thrust reverser cycling test is used to detect any operational faults of the thrust reverser system including the electrical interfaces, mechanical linkage, and translating sleeve. This is done without an engine start.
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CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
MAINT MENU EEC1-A-THRUST REVERSER CYCLING TEST
1/8
*** WARNING *** THRUST REVERSER WILL DEPLOY AND STOW DURING THIS TEST ENSURE THERE ARE NO PERSONNEL NEAR THRUST REVERSER
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- FADEC IS POWERED 3- THRUST REVERSER HYDRAULIC PRESSURE IS AVAILABLE 4- THRUST REVERSER IS NOT INHIBITED 5- THROTTLE LEVER IS SET TO FWD IDLE 6- PRESS "CONTINUE" TO GO TO NEXT PAGE
* NOTE *
YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7321_026
TEST WILL TIMEOUT AFTER 5 MINUTES INACTIVITY
Figure 33: Thrust Reverser Cycling Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-69
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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78-70
CS130_78_ILB.fm
ATA 79 - Oil
BD-500-1A10 BD-500-1A11
79 - Oil
Table of Contents 79-11 Storage .........................................................................79-2 General Description ..........................................................79-2 Practical Aspects .............................................................79-4 Oil Servicing ................................................................79-4 79-20 Distribution....................................................................79-6 General Description .........................................................79-6 Component Location ........................................................79-8 Lubrication and Scavenge Oil Pump............................79-8 Oil Control Module .......................................................79-8 Deoiler........................................................................79-10 Fuel/Oil Heat Exchanger Bypass Valve .....................79-10 Fuel/Oil Heat Exchanger............................................79-10 Variable Frequency Generator Oil/Oil Heat Exchanger .............................79-10 Air/Oil Heat Exchanger ..............................................79-10 Component Information ..................................................79-12 Lubrication and Scavenge Oil Pump .........................79-12 Oil Control Module .....................................................79-14 Deoiler........................................................................79-16 Main Oil Filter ............................................................79-18 Fan Drive Gear System Oil Pump..............................79-20 Detailed Component Information ....................................79-22 Journal Oil Shuttle Valve............................................79-22 Variable Oil Reduction Valve .....................................79-22 Detailed Description ........................................................79-24 Oil System Control .....................................................79-26
Detailed Component Information ................................... 79-34 Oil Debris Monitor ..................................................... 79-34 Controls and Indications ................................................ 79-36 EICAS ....................................................................... 79-36 Status ........................................................................ 79-36 Detailed Description ...................................................... 79-38 Oil Pressure Warning Limits...................................... 79-40 Monitoring and Tests ..................................................... 79-42 CAS Messages ......................................................... 79-43
79-30 Indicating ....................................................................79-30 General Description .......................................................79-30 Component Location ......................................................79-32 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-i
CS130_M5_79_OilTOC.fm
79 - Oil
List of Figures Figure 1: Oil Storage ............................................................79-3 Figure 2: Oil Servicing ..........................................................79-5 Figure 3: Oil Distribution .......................................................79-7 Figure 4: Component Location Right Side............................79-9 Figure 5: Component Location Left Side ............................79-11 Figure 6: Lubrication and Scavenge Oil Pump ...................79-13 Figure 7: Oil Control Module...............................................79-15 Figure 8: Deoiler .................................................................79-17 Figure 9: Main Oil Filter ......................................................79-19 Figure 10: Fan Drive Gear System Oil Pump .......................79-21 Figure 11: Journal Oil Shuttle Valve and Variable Oil Reduction Valve...............................79-23 Figure 12: Distribution Schematic.........................................79-25 Figure 13: Oil System Control ..............................................79-27 Figure 14: Fan Drive Gear System Operation ......................79-29 Figure 15: Oil Indicating........................................................79-31 Figure 16: Oil System Indicating Components .....................79-33 Figure 17: Oil Debris Monitor................................................79-35 Figure 18: Oil Indications......................................................79-37 Figure 19: Oil Indicating Detailed Description ......................79-39 Figure 20: Oil Pressure Warning Limits................................79-41
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79-ii
CS130_M5_79_OilLOF.fm
79 - Oil
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TECHNICAL TRAINING MANUAL
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79-iii
CS130_M5_79_OilLOF.fm
OIL - CHAPTER BREAKDOWN OIL CHAPTER BREAKDOWN
Storage
1
Distribution
2
Monitoring and Tests
3
79 - Oil 79-11 Storage
79-11 STORAGE GENERAL DESCRIPTION A 27.3 L (28.8 qt) aluminum oil tank stores the oil and supplies it to the engine. The tank is wrapped in a stainless steel heatshield, held together by springs. The oil tank has an internal swirl-type deaerator that removes air from the returning oil. A pressure-regulating valve (PRV) maintains 12 psi head pressure in the tank. Excess pressure is vented to the deoiler in the main gearbox. The oil level can be monitored using a sight glass. A fill port has a hinged cap for oil servicing. The fill port strainer prevents foreign objects from entering the tank during servicing. In the event that the filler cap is incorrectly installed, a flapper valve prevents rapid loss of oil. A scupper drain collects oil spilt during servicing, and sends it to the drain mast. Oil can be drained through a drain plug in the bottom of the tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-2
CS130_79_L2_Storage_General.fm
79 - Oil 79-11 Storage
Oil Fill Port
Sight Glass
A
Pressurization Valve Scavenge Return Spring-Loaded Lock To Deoiler
Hinged Cap
To Lubrication and Scavenge Oil Pump
Flapper Valve A OIL TANK
B
B OIL FILL PORT
Scupper Drain
CS1_CS3_7911_002
Drain Plug
Figure 1: Oil Storage Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-3
CS130_79_L2_Storage_General.fm
79 - Oil 79-11 Storage
PRACTICAL ASPECTS
Continue to hold the flapper valve open. Add the recommended oil into the filler neck until no more oil can be added without overflow into the scupper drain.
OIL SERVICING The oil level can be checked and serviced by opening an access door on the right side of the core cowl fixed structure. The tank is equipped with a flapper valve to ensure the oil is not lost, should the cap remain open during flight. The sight glass is marked in both liters and quarts, and gauge markings indicate how much oil may be added.
WARNING A MINIMUM WAIT OF 5 MINUTES IS REQUIRED TO MAKE SURE THE OIL SYSTEM IS NOT PRESSURIZED PRIOR TO SERVICING THE OIL. FAILURE IN FOLLOWING THIS WARNING MAY RESULT IN PERSONAL INJURY.
Oil quantities are as follows: •
Fully serviced 21.8 L (23 qts)
•
Minimum 9.3 L (9.8 qts)
CAUTION
If the oil level is checked between 5 minutes and 60 minutes after engine shutdown, add oil to bring the oil level to the FULL mark. If the oil level is checked between 1 hour and 10 hours after engine shutdown, add oil only if the level is below the 3 L (3qt) mark on the sight glass. Do not add more than 3 L (3.1 qt). If the oil level is still not at the 3 L (3qt) mark, the engine must be run at idle until the oil is at the correct operating temperature. Shut down the engine and check the oil level. If required, add oil to bring the oil level to the FULL mark. If the oil level is checked more than 10 hours after engine shutdown, dry motor the engine until the engine oil pressure is stable. Check the oil quantity on the STATUS synoptic page. If the oil level is less than 15.1 L (16.0 US qts), add oil until the oil level is at the 3 L (3 qts) mark on the sight glass. Run the engine at ground idle until it is at operating temperature, then shutdown the engine. If required, add oil to bring the oil level to the FULL mark. To service the oil, lift the hinged cap and look into the oil tank to ensure the flapper is not stuck in a closed position. If the flapper is stuck, insert a small screwdriver through one of the holes in the oil tank inlet screen and open the oil tank flapper valve. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
1. Check the oil level between 5 minutes and 1 hour after engine shutdown. Outside of this range, the oil level sight glass indication is not accurate. 2. Do not motor the engine before checking the oil level during the 1 hour to 10 hours after engine shutdown interval. Too much oil can be added and cause damage to the engine. 3. Do not damage the flapper valve when using a screwdriver to open it. 4. Use only engine oil specified in the service bulletin. The mixing of different brands of approved oils is not recommended, but is permitted within the limits specified in the service bulletin. The use of unapproved types or brands of oils is not permitted, and may cause damage to the engine. 5. Make sure the oil tank is completely closed. Improper closure and locking of the handle may result in oil exiting the tank, resulting in an inflight shutdown.
TECHNICAL TRAINING MANUAL
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79-4
CS130_79_L2_Storage_PracticalAspects.fm
79 - Oil 79-11 Storage
Spring-Loaded Lock Hinged Cap
A B
Flapper Valve
B OIL LEVEL SIGHT GLASS A
OIL FILL PORT
CS1_CS3_7911_003
OIL TANK
Figure 2: Oil Servicing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-5
CS130_79_L2_Storage_PracticalAspects.fm
79 - Oil 79-20 Distribution
79-20 DISTRIBUTION GENERAL DESCRIPTION The main function of the engine oil system is to cool and lubricate the bearings, gearboxes, shafts and splines throughout the engine, and provide system status and fault indications for display on the engine indication and crew alerting system (EICAS). It also provides a thermal management function, where heat exchangers are used to cool the oil and heat the fuel. The system consists of the oil tank, pressure pump, oil control module (OCM), filters, heat exchangers, and associated plumbing. The plumbing connections on the engine core have been minimized by the use of the OCM, which is the central distribution point in the oil system. The dual-element main oil filter, installed on the OCM, has a primary filter and a primary filter bypass valve, which bypasses flow around the primary filter element in case of filter clogging. A secondary filter still continues to provide filtering in the event of a bypass. In addition to supplying oil pressure to the engine gearbox and bearings, the lubrication and scavenge oil pump (LSOP) also removes oil from the bearing cavities and gearboxes, and returns it to the tank. Pressure buildup in the compressor intermediate case causes the no. 3 bearing compartment and angle gearbox to experience increased pressure that must be released. The released air pressure, mixed with oil vapor, is called breather air. An external tube sends no. 3 bearing breather air to a deoiler unit in the main gearbox. The deoiler removes the oil droplets from the air. This air is sent overboard through the deoiler vent duct. Additional breather air comes from the oil tank. Oil tank pressure is released through the tank pressure valve when the tank pressure reaches the maximum limit The fan drive gear system (FDGS) has a dual-stage fan-driven oil pump that works with the journal oil shuttle valve (JOSV) to form an integrated Copyright © Bombardier Inc. CS130-21.07-06-00-121418
oil system, supplying oil to the FDGS journal bearings during normal, windmill, and negative-G conditions. A variable oil reduction valve (VORV) supplies pressurized oil from the engine oil system during normal operation. The VORV controls the amount of oil supplied based on engine power settings. The front bearing compartment contains two sumps that return oil back to the main oil tank from the components in the front bearing compartment. The sump, at the bottom of the front bearing compartment, also acts to supply oil to the journal bearings during windmill conditions. In normal operation the sump oil is scavenged and returned to the oil tank by the fan-driven oil pump. A secondary sump collects oil from the FDGS via an oil collection gutter. This secondary sump in the front compartment supplies oil during negative-G events. In normal operation, the fan-driven oil pump scavenges the oil and returns it to the oil tank. The thermal management system, consisting of three heat exchangers, ensures that engine oil and fuel temperatures are maintained within limits. The air/oil heat exchanger (AOHX) uses the fan air to cool the engine oil. If the AOHX becomes clogged, the pressure-relief valve (PRV) diverts oil directly to the variable frequency generator (VFG) oil/oil cooler. The fuel/oil heat exchanger (FOHX) is used to decrease oil temperature, and increase fuel temperature by transferring heat from the engine oil to the engine fuel. The FOHX has integral pressure-relief valves for both the fuel and oil sides of the heat exchanger. The fuel/oil heat exchanger bypass valve (FOHXBV) is a modulating valve used to control oil flow between the FOHX and AOHX based on fuel temperature. The variable frequency generator oil/oil heat exchanger (VFGOOHX) uses cooled engine oil to remove heat from the oil used by the VFG. An oil debris monitor and multiple chip collectors located on the LSOP provide system health monitoring.
TECHNICAL TRAINING MANUAL
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79-6
CS130_79_L2_Distribution_General.fm
79 - Oil 79-20 Distribution
LEGEND
OIL TANK
AGB Deoiler Overboard Breather No. 3 Bearing Breather
MAIN ENGINE BEARINGS 1,1.5,2,3,4,5,6 OCM
LUBRICATION AND SCAVENGE OIL PUMP
ODM
FDGS
Main Oil Filter
Fan Air
FDG VORV JOSV
AIR/OIL HEAT EXCHANGER
Fan Air Overboard
MGB
FUEL/OIL HEAT EXCHANGER BYPASS VALVE To Main Pump
To VFG
FUEL/OIL HEAT EXCHANGER OIL/OIL HEAT EXCHANGER
Journal Bearings
FDOP
From Boost Pump
Auxiliary Pump Stage
SECONDARY SUMP
Windmil Pump Stage
SUMP
From VFG
CS1_CS3_7920_003
AGB FDG FDGS FDOP JOSV MGB OCM ODM VFG VORV
FDGS Circuit Breather Air Fuel Pressure Scavenge VFG Oil System Supply Angle Gear Box Fan Drive Gear Fan Drive Gear System Fan-Driven Oil Pump Journal Oil Shutoff Valve Main Gear Box Oil Control Module Oil Debris Monitor Variable Frequency Generator Variable Oil Reduction Valve
Figure 3: Oil Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-7
CS130_79_L2_Distribution_General.fm
79 - Oil 79-20 Distribution
COMPONENT LOCATION The oil distribution system includes the following components: •
Lubrication and scavenge oil pump (LSOP)
•
Oil control module (OCM)
•
Deoiler
•
Main oil filter
•
Active oil damper valve (AODV)
•
Variable oil reduction valve (VORV)/journal oil shuttle valve (JOSV)
•
Fuel/oil heat exchanger bypass valve (FOHXBV) (Refer to figure 5)
•
Fuel/oil heat exchanger (FOHX) (Refer to figure 5)
•
Variable frequency generator oil/oil heat exchanger (VFGOOHX) (Refer to figure 5)
•
Air/oil heat exchanger (AOHX) (Refer to figure 5)
LUBRICATION AND SCAVENGE OIL PUMP Lubrication and scavenge oil pump (LSOP) is installed on the right front side of the main gearbox at the 5 o’clock position. OIL CONTROL MODULE Oil control module is mounted to the right side of the main gearbox.
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79-8
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
Active Oil Damper Valve
Oil Filter
Variable Oil Reduction Valve Lubrication and Scavenge Oil Pump
Journal Oil Shuttle Valve OIL CONTROL MODULE
CS1_CS3_7920_009 920_009
Deoiler
Figure 4: Component Location Right Side Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-9
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
DEOILER The deoiler is installed on the right side of the main gearbox. FUEL/OIL HEAT EXCHANGER BYPASS VALVE The fuel/oil heat exchanger bypass valve installed on the front side of the core engine at the 12 o’clock position. FUEL/OIL HEAT EXCHANGER The fuel/oil heat exchanger is installed on the left side of the core at the 11 o’clock position. VARIABLE FREQUENCY GENERATOR OIL/OIL HEAT EXCHANGER The variable frequency generator oil/oil heat exchanger is installed on the left side of the core engine at approximately the 10 o’clock position. AIR/OIL HEAT EXCHANGER The air/oil heat exchanger is installed on the left side of the core engine at the 10 o’clock position.
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79-10
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
A C
D
A
B
FUEL/OIL HEAT EXCHANGER
C
VFG OIL/OIL HEAT EXCHANGER
FUEL/OIL HEAT EXCHANGER BYPASS VALVE
D
AIR/OIL HEAT EXCHANGER
CS1_CS3_7920_008
B
Figure 5: Component Location Left Side Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-11
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
COMPONENT INFORMATION LUBRICATION AND SCAVENGE OIL PUMP The lubrication and scavenge oil pump (LSOP) provides pressure for lubricating components, and a scavenge function to draw the oil back to the oil control module (OCM) and tank for reuse. The LSOP has seven stages. Six of the stages scavenge oil from the bearing compartments and gearboxes. The seventh stage pressurizes the oil, and sends it to the oil control manifold where it then goes directly to the main oil filter. Six magnetic chip collectors in each of the return lines catch metallic particles. The collectors catch ferrous metal particles in the scavenge oil, which can then be used to diagnose system problems. The chip collectors have quick removal type fittings for easy maintenance. Chip collectors are labeled according to the scavenge stage the oil is collected from: •
Front bearing compartment (FBC), including the fan drive gear system (FDGS) and bearings no. 1, no. 1.5, and no. 2
•
No. 3 bearing compartment (BC3)
•
No. 4 bearing compartment (BC4)
•
No. 5 and no.6 bearing compartments (BC5)
•
Main gearbox (MGB)
•
Angle gearbox (AGB)
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79-12
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
LEGEND Accessory Gear Box Bearing Compartment 3 Bearing Compartment 4 Bearing Compartment 5 Forward Bearing Compartment Main Gear Box
Housing
Chip Collector
CS1_CS3_7920_004
AGB BC3 BC4 BC5 FBC MGB
Figure 6: Lubrication and Scavenge Oil Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-13
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
OIL CONTROL MODULE The oil control module (OCM) is used to mount related oil system components. Various line replaceable units (LRUs) are centered around the body of the OCM to simplify maintenance. The OCM provides various oil passages between components. Oil flows into and through this module, using external oil tubes, and OCM housing passages.The oil metering plugs control the flow of oil and can be replaced to allow system pressures and engine flow to be adjusted if required. The following distribution system components are mounted on the OCM: •
Main oil filter (MOF)
•
Variable oil reduction valve (VORV)
•
Journal oil shuttle valve (JOSV)
•
Active oil damper valve (AODV)
•
Oil debris monitor (ODM)
•
Oil metering plugs
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79-14
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Active Oil Damper Valve (AODV)
Main Oil Filter
Oil Debris Monitor
Variable Oil Reduction Valve/Journal Oil Shuttle Valve (VORV/JOSV)
Oil Metering Plugs
CS1_CS3_7920_005
Oil Metering Plug
Figure 7: Oil Control Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-15
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
DEOILER During engine operation, sealing air flows into the bearing compartments. The sealing air must be vented to allow a continuous flow of air. The vented sealing air is referred to as breather air. The engine oil breather system removes air from the bearing compartments, separates breather air from the oil, and vents the air overboard. The deoiler receives air from the main gearbox, the no. 3 bearing compartment, and the deaerator in the oil tank. The deoiler separates the oil from the air. The oil returns to the gearbox for scavenging, and the air is vented overboard through the deoiler vent duct. The line replaceable deoiler drive oil seal creates a seal between the deoiler shaft and the main gearbox.
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79-16
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Deoiler Vent Duct
MAIN GEARBOX
CS1_CS3_7920_010
Deoiler
Figure 8: Deoiler Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-17
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
MAIN OIL FILTER The main oil filter assembly is a component of the OCM. It removes contaminants from the oil as it exits the LSOP. The main oil filter assembly consists of the main oil filter housing, machined into the OCM, a main oil filter element, and a main oil filter bypass valve. The main oil filter is covered by a thermal blanket for heat protection. The main oil filter element is a reverse flow type, meaning the oil flow path is from the inner diameter of the filter to the outside. If the main oil filter clogs, the oil is bypassed into a secondary filter within the main filter assembly. The main oil filter housing has two jackscrew lands to facilitate removal of the cover. The housing has a built-in drain and vent plugs to allow drainage before the cover is removed.
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79-18
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Vent Plug
Secondary Filter
Thermal Blanket Cover Oil Filter Cover
Drain Plug
Primary Filter Jackscrew Lands Thermal Blanket
CS1_CS3_7920_006
Oil Filter Bypass
Figure 9: Main Oil Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-19
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
FAN DRIVE GEAR SYSTEM OIL PUMP The dual stage pump continuously draws oil from a dedicated auxiliary reservoir and compartment sump located in the front bearing compartment. The pump is installed in the no. 1 and no. 1.5 bearing support, and provides the journal bearings with an alternate oil supply during engine windmilling and negative-G conditions. It is turned by the fan drive shaft and rotates at fan speed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-20
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Auxiliary Reservoir
Fan Oil Pump
CS1_CS3_7920_012
No. 1 and No. 1.5 Bearing Support
Figure 10: Fan Drive Gear System Oil Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-21
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
DETAILED COMPONENT INFORMATION JOURNAL OIL SHUTTLE VALVE The journal oil shuttle valve (JOSV) is a mechanical, two position device that directs oil flow from the main oil or fan drive gear system oil supply to the journal bearings. The JOSV is controlled by comparing the main oil pressure against the gearbox vent pressure. When oil pressure is normal, the primary oil goes to the journal bearings. If oil pressure decreases, the JOSV sends oil from the fan drive gear system to the journal bearings. VARIABLE OIL REDUCTION VALVE The variable oil reduction valve (VORV) diverts the supply of oil from the FDGS to the oil tank. The valve is controlled by the electronic engine control (EEC) through an electrohydraulic servovalve (EHSV). The EEC receives valve position feedback from a linear variable differential transformer (LVDT). Maximum oil flow to lubricate the gear faces of the FDGS is required at takeoff only. At cruise, oil flow is reduced and sent back to the oil tank. Less oil flowing to the gears reduces the oil heat load and increases fan drive gearbox efficiency. When in the failsafe position the valve reverts to maximum oil flow.
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79-22
CS130_79_L3_Distribution_DetailedComponentInformation.fm
79 - Oil 79-20 Distribution
JOURNAL OIL SHUTTLE VALVE
From Windmill Pump
Gearbox Vent
From Main Oil Supply
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
JOURNAL OIL SHUTTLE VALVE
From Windmill Pump
LVDT
LVDT
T/M
T/M
Gearbox Vent
To Journal Bearings
ReturnTo Oil Tank
To Fan Drive Gear System
To Journal Bearings
ReturnTo Oil Tank
JOSV NORMAL FLOW
JOURNAL OIL SHUTTLE VALVE
Gearbox Vent
From Main Oil Supply JOURNAL OIL SHUTTLE VALVE
From Windmill Pump
From Main Oil Supply
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
LVDT
LVDT
T/M
T/M
Gearbox Vent
To Journal Bearings
ReturnTo Oil Tank
To Fan Drive Gear System VORV TAKEOFF
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
From Windmill Pump
From Main Oil Supply
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
To Fan Drive Gear System JOSV LOW MAIN OIL PRESSURE
To Journal Bearings
ReturnTo Oil Tank
To Fan Drive Gear System VORV CRUISE
LVDT T/M
Fan Drive Gear System Oil Pressure Return Vent Liner Variable Differential Transformer Torque Motor
NOTE JOSV shown in normal position. VORV shown in low flow position.
CS1_CS3_7921_001
LEGEND
Figure 11: Journal Oil Shuttle Valve and Variable Oil Reduction Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-23
CS130_79_L3_Distribution_DetailedComponentInformation.fm
79 - Oil 79-20 Distribution
DETAILED DESCRIPTION The oil leaves the tank through an oil strainer, passes through an oil pressure element of the LSOP, then to the OCM where it is filtered by the main oil filter and distributed. If the filter begins clogging, an OIL FILTER IMPENDING BYPASS info message is shown. The filter bypasses at 55 psid as indicated by a L(R) OIL FILTER caution message and the secondary oil filter maintains a sufficient level of filtering. The no. 3 damper supply is fed just downstream of the filter and the remainder of the oil is then split between cooled and uncooled paths. The cooled path is delivered to the heat exchangers where it is cooled and returned to the OCM. From there, some of the cooled oil is supplied to the FDGS journal bearings, and the remainder of the cooled oil mixes with the uncooled oil and is distributed to lubricate the engine main shaft bearings and seals, FDGS gears, and both the accessory and angle gearboxes. After these components have been lubricated, the oil is scavenged by the LSOP and pumped back through the OCM to the oil tank, where it is deaerated. The LSOP has six scavenge pump elements that pull oil from the: •
Front bearing compartment servicing the FDGS and bearings no. 1, no. 1.5, and no. 2
•
No. 3 bearing compartment
•
No. 4 bearing compartment
•
No. 5 and no. 6 bearing compartment
•
Main gearbox
•
Angle gearbox
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The oil drains into the tank, while the breather air/oil mixture leaves the tank and enters the gearbox. A rotating deoiler separates the air/oil mixture, allowing oil to be scavenged by the gearbox scavenge pump, and pass air to the overboard vent. Large particles could enter the oil supply beyond the main oil filter on the oil control module. Last chance oil strainers located in the supply lines prevent these particles from entering bearing compartments and clogging oil nozzles. The last chance strainers are located: •
FDGS gears
•
No. 1, no. 1.5, and no. 2 bearing compartments
•
No. 3 bearing compartment
•
No. 4 bearing compartment
•
No. 5 and no. 6 bearing compartments
•
Angle gear box
Heating and cooling of the engine fuel and oil is accomplished by modulating the flow of oil through a fuel/oil heat exchanger (FOHX) and air/oil heat exchanger (AOHX). The proportion of oil that flows through the two types of heat exchanger are controlled by the electronic engine control (EEC) and the fuel/oil heat exchanger bypass valve (FOHXBV). The FOHXBV is controlled based on the fuel temperature, ensuring that the engine oil and engine fuel temperatures are maintained within limits.
TECHNICAL TRAINING MANUAL
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79-24
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
OIL CONTROL MODULE VORV MOP
JOSV
No. 4
Journal Bearings
LUBRICATION AND SCAVENGE OIL PUMP
MOT
Pressure Valve De-aerator
Deoiler Overboard
A O P
MGB VFG
Oil Level Sensor
ODM
No. 3
No. 6
MOF
Active Oil Damper Valve
No. No. Gear Teeth No. 1 1.5 2 Auxiliary Dual-Stage Tank Windmill Pump
No. 5
OIL TANK
PMAG
FDGS AUXILIARY CIRCUIT LEGEND
ANGLE GEARBOX
FOHX FOHX MANIFOLD AIR/OIL HEAT EXCHANGER VFG AIR/OIL HEAT EXCHANGER
OIL/OIL HEAT EXCHANGER
FUEL/OIL MANIFOLD
MAIN GEARBOX
AOP FDGS FOHX JOSV MOF MOP MOT
Breather Air ODM VFG Scavenge PMAG VFG Oil Supply VFG Pressure Scavenge VORV Fan Drive Gear System Auxiliary Oil Pressure Fan Drive Gear System Fuel/Oil Heat Exchanger Journal Oil Shuttle Valve Main Oil Filter Main Oil Pressure Main Oil Temperature
Oil Debris Monitor Permanent Magnet Alternator Generator Variable Frequency Generator Variable Oil Reduction Valve Chip Collector Last Chance Screen Valve
CS1_CS3_7930_005
FOHX Bypass Valve
Figure 12: Distribution Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-25
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
OIL SYSTEM CONTROL
Variable Oil Reduction Valve
The EEC controls the following oil system components:
The VORV is an electromechanical valve that diverts oil from the FDGS to the oil tank. The EEC controls the dual-channel torque motor driven VORV, while a linear variable differential transformer (LVDT) provides position feedback. Oil flow to lubricate the gear faces is only required at takeoff. At cruise, less oil flow is required, so the VORV diverts some of the oil back to the oil tank, to maximize efficiency during cruise conditions. The amount of oil required is a function of the load placed on the FDGS. The amount of oil is scheduled as a function of horsepower extracted by the fan.
Active Oil Damper Valve The no. 3 bearing active oil damper valve (AODV) controls the oil supply to the no. 3 bearing damper. The oil supply is controlled to: •
Optimize no.3 bearing loads during all phases of operation
•
Limit high-pressure spool vibration
•
Provide bowed rotor protection at sub-idle operation
The solenoid operated valve is controlled by either channel of the EEC based on N2 rpm. The valve is open during start up to N2 idle speed (62%). The valve is energized closed up to 85% N2, then reopens. Fuel/Oil Heat Exchanger Bypass Valve The FOHXBV is an electromechanical modulating valve that controls the amount of oil going to the FOHX and/or the AOHX.
The fan thrust is calculated based on shaft speed, air pressure, and temperature entering the fan. Due to the similarity in the thrust and P3 burner pressure sensor characteristic response over the flight envelope, the VORV is scheduled as a function of P3 for above idle operation. If the VORV is stuck in the low flow setting, then inadequate lubrication at high engine power settings may cause degradation of the FDGS gears. If stuck in the high flow setting, the oil temperature reads high at low power settings.
The EEC receives the fuel temperature sensor signal and varies the amount of oil going to each heat exchanger, based on the fuel temperature sensor readings. The FOHXBV supplies a minimum of 25%, and a maximum of approximately 100% of the oil flow to the AOHX. Under normal operating conditions, the FOHXBV splits oil flow as follows: •
75% is supplied to the fuel/oil heat exchanger
•
25% is suppled to the air/oil heat exchanger
The FOHXBV increases oil flow through the FOHX if the fuel temperature is too low. The increased oil flow through the FOHX raises the fuel temperature.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-26
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
From VFG EEC
To VFG Air/Oil Heat Exchanger
VFG OIL/OIL HEAT EXCHANGER
AIR/OIL HEAT EXCHANGER
Channel A
FOHXBV T/M
OCM
To Oil System
LVDT
Channel B
To No. 3 Bearing
AODV
To FDGS
Fuel FUEL/OIL HEAT EXCHANGER
To Oil Tank
FUEL TEMPERATURE SENSOR
T/M
VORV
From Oil Pump
P3 BURNER PRESSURE SENSOR
LEGEND
CS1_CS3_7930_006
AODV EEC FDGS FOHXBV LVDT OCM T/M VFG VORV
Fuel Oil VFG Oil Active Oil Damper Valve Electronic Engine Control Fan Drive Gear System Fuel/Oil Heat Exchanger Bypass Valve Linear Variable Differential Transformer Oil Control Module Torque Motor Variable Frequency Generator Variable Oil Reduction Valve CAN BUS
Figure 13: Oil System Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-27
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
Fan Drive Gear System Oil Circuit The fan drive gear system (FDGS) journal bearings are lubricated with pressurized oil from a manifold at the rear of the FDGS. The FDGS oil supply is provided by the variable oil reduction valve (VORV) and the journal oil shuttle valve (JOSV). Both valves ensure the fan drive gear system (FDGS) receives the correct amount of oil when required. During normal operating conditions, the main oil pump pressurizes the oil system through the VORV and the JOSV returns the sump oil to the oil tank. A dual-stage fan-driven pump acts as a supplemental scavenge pump, drawing oil from the sump and returning it to the oil tank. During engine windmilling in flight, on the ground, or during negative-G conditions, the main oil pressure decreases. The journal oil shuttle valve moves to isolate the journal bearing oil flow from the main oil supply, allowing the fan-driven oil pump to supply oil from the sump to the journal bearings. In a negative-G condition, the fan-driven pump draws oil from the secondary sump to supply oil to the journal bearings. The operation of the journal oil shuttle valve and fan-driven oil pump are verified on every flight via the auxiliary oil pressure sensor. This detects both positive oil flow from the fan-driven oil pump, and shuttle valve movement from the main oil pump supply position to the fan-driven pump oil supply position. Rotation of the fan for maintenance does not impact the oil system components. A failure of the fan-driven oil pump or JOSV is detected by the auxiliary oil pressure sensor and indicated at the next engine start.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-28
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
NORMAL OPERATION Journal Bearing Oil Filter
NEGATIVE-G OR WINDMILL OPERATION Journal Bearing Oil Filter
VORV/JOSV
NO. 1/1.5 BEARING SUPPORT
VORV/JOSV
NO. 1/1.5 BEARING SUPPORT
Oil Manifold
Oil Manifold OIL TANK
OIL TANK
FDGS/ Journal Bearings
FDGS/ Journal Bearings
Gutter
Gutter
Lubrication Pump
Secondary Sump
Scavenge Pump
Fan Oil Pump
Secondary Sump
Lubrication Pump
Scavenge Pump
Fan Oil Pump
LEGEND
FDGS VORV/JOSV
Sump
FDGS Pressure Scavenge Supply Fan Drive Gear System Variable Oil Reduction Valve/Journal Oil Shuttle Valve Auxiliary Oil Pressure Sensor
CS1_CS3_7930_015
Sump
Figure 14: Fan Drive Gear System Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-29
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-30 Indicating
79-30 INDICATING GENERAL DESCRIPTION The oil indicating system provides oil temperature, pressure, and quantity information for display in the flight deck. All of the sensors and probes are line replaceable units (LRUs). The oil quantity sensor is mounted on top of the oil tank. The oil quantity sensor provides the quantity of engine oil in the oil tank. The dual-channel oil pressure sensor measures the pressure output of the lubrication and scavenge oil pump (LSOP). The dual-channel oil temperature probe measures the temperature of the scavenge oil returning to the oil tank. The dual-channel oil filter differential pressure sensor monitors the differential pressure across the main oil filter, and provides an indication of an impending filter bypass. The oil debris monitor (ODM) detects both ferrous, and non-ferrous debris in the oil system. The presence of metallic particles is detected by a sense coil in the ODM. This produces a signal recorded by the prognostics and health management unit (PHMU), which classifies the particles according to size, and whether it is ferrous or non-ferrous. The PHMU records the data for engine health monitoring purposes. A dual-element auxiliary oil pressure sensor monitors oil pressure in the fan drive gear system (FDGS). The oil circuit detects failures of the dual-stage pump and the journal oil shuttle valve (JOSV). The sensors and probes supply information to the electronic engine control (EEC). The EEC provides the information for display on the engine indication and crew alerting system (EICAS), the STATUS synoptic page, the onboard maintenance system (OMS), and the health management unit (HMU).
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79-30
CS130_79_L2_Indicating_General.fm
79 - Oil 79-30 Indicating
Oil Quantity Sensor
TO 90.8
90.8
18.6
18.6
N1
516 OIL TANK
EGT ELECTRONIC ENGINE CONTROL
Oil Filter Differential Pressure Sensor
522
62.1 220 103 113
Oil Pressure Sensor
N2 FF (KPH) OIL TEMP OIL PRESS
62.0 220 103 113
EICAS Oil Temperature Sensor
OIL CONTROL MODULE
Oil Debris Monitor
Excitation
ENGINE
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L)
103 113 21.8
SYNOPTIC PAGE – STATUS LEGEND ARINC 429 CAN BUS Analog
OMS HMU
CS1_CS3_7930_016
AUXILIARY OIL PRESSURE SENSOR
Figure 15: Oil Indicating Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-31
CS130_79_L2_Indicating_General.fm
79 - Oil 79-30 Indicating
COMPONENT LOCATION The following oil indicating component is located on the oil tank: •
Oil quantity sensor
The following components are located on the oil control module (OCM): •
Oil filter differential pressure sensor
•
Oil debris monitor
•
Oil pressure sensor
•
Oil temperature sensor
The following auxiliary oil system component is located at the 5 o’clock position on the fan intermediate case. •
Auxiliary oil pressure sensor
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CS130_79_L2_Indicating_ComponentLocation.fm
79 - Oil 79-30 Indicating
Oil Quantity Sensor Oil Filter Differential Pressure Sensor
A
Oil Pressure Sensor
B D
Oil Temperature Sensor A
OIL TANK
B
OIL CONTROL MODULE
C
OIL DEBRIS MONITOR
D
AUXILIARY OIL PRESSURE SENSOR
CS1_CS3_7930_008
C
Figure 16: Oil System Indicating Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-33
CS130_79_L2_Indicating_ComponentLocation.fm
79 - Oil 79-30 Indicating
DETAILED COMPONENT INFORMATION OIL DEBRIS MONITOR The ODM sensor core consists of a magnetic coil assembly with two field coils and one sense coil. The coil assembly receives an excitation signal from the PHMU. Two equal and opposing magnetic fields are generated, which balance in the center. Passage of a metallic particle disrupts the balance of the fields seen by the sense coil, and results in a characteristic signal. The signal resembles a sine wave: •
Amplitude is proportional to particle size
•
Phase of ferrous particle signal versus non-ferrous particle signal is offset by 90°
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-34
CS130_79_L3_Indicating_DetailedComponentInformation.fm
79 - Oil 79-30 Indicating
Field Coils
Sense Coil
Coil Assembly
Particle
Signal (volt)
Signal ~ Particle Size
LEGEND Ferrous Non-Ferrous
Excitation Signal Oil Passage Tube
Coils
OIL DEBRIS MONITOR CROSS – SECTION VIEW
CS1_CS3_7930_002
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
Figure 17: Oil Debris Monitor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-35
CS130_79_L3_Indicating_DetailedComponentInformation.fm
79 - Oil 79-30 Indicating
CONTROLS AND INDICATIONS EICAS The oil temperature and pressure are displayed on the EICAS page. STATUS The oil temperature, pressure, and quantity are displayed on the STATUS synoptic page.
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TECHNICAL TRAINING MANUAL
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79-36
CS130_79_L2_Indicating_ControlsIndications.fm
79 - Oil 79-30 Indicating
TO 90.8
90.8
18.6
18.6
OIL TEMPERATURE Symbol
N1
516
Description
103
Oil temperature normal range
158
Oil temperature above high oil temperature yellow line
183
Oil temperature above high oil temperature red line threshold
35
522
èèè
Oil temperature below low oil temperature threshold Oil temperature invalid
EGT N2 FF (KPH) OIL TEMP OIL PRESS
62.0 220 103 113
OIL PRESSURE Symbol
81
AIR IR
DOOR OR
ELEC C
FLT CTR CTRL TRL L
FUEL FUE EL
HYD YD
AVIONIC IONIC
INFO
CB
TAT SAT
Oil pressure above high oil pressure threshold
20
Oil pressure below low oil pressure threshold
èèè
Oil pressure invalid
15°C 15°C
OIL QUANTITY Symbol
ENGINE
103 113 21.8
Oil pressure normal range
203
EICAS
STATUS
Description
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (QTS)
19.2
103 113 21.8
Description Normal
2.0
Below threshold
èèè
Invalid
STATUS PAGE
CS1_CS3_7930_018
62.1 220 103 113
Figure 18: Oil Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-37
CS130_79_L2_Indicating_ControlsIndications.fm
79 - Oil 79-30 Indicating
DETAILED DESCRIPTION The oil quantity sensor sends an oil quantity signal to channel B of the EEC. When the main oil tank quantity is at or below 9.45 L (10 qt), a cyan L (R) ENG OIL LO QTY message is displayed on the EICAS page. The oil filter differential pressure sensor signal is sent to both channels of the EEC. When the differential pressure across the filter is more than 35 psid at rated flow conditions, an OIL FILTER IMPENDING BYPASS INFO message is displayed. When the differential pressure across the oil filter element is more than 55 psid, the filter bypass valve opens and an L (R) ENG OIL FILTER caution message is displayed.
An AUX OIL PRESS MON INOP INFO message is displayed when the auxiliary oil pressure sensor fails. Table 1: Oil System Limits OIL SYSTEM LIMITS PARAMETER Oil Quantity Oil Filter Bypass
LIMIT
WARNING
< 9.45 L (10 qt)
L(R) ENG OIL LO QTY advisory message.
35 psid
Oil filter impending bypass INFO message.
55 psid
L(R) ENG OIL FILTER caution message.
The oil pressure sensor sends signals to both channels of the EEC. The oil temperature probe sends oil temperature information to both channels of the EEC.
Oil Temperature
An ENG OIL LO TEMP caution message is displayed if the throttles are advanced beyond 50% N1 before the oil temperature has reached 49°C (120°F).
L(R) ENG EXCEEDANCE caution message.
Minimum oil temperature for starting is - 40° C (40° F). Operation at maximum oil temperature of 174° C (345° F) cannot exceed 20 minutes. Oil Pressure
The operation of the JOSV and the fan-driven oil pump are verified by the auxiliary oil pressure sensor, which detects positive oil flow from the fan-driven oil pump. It also detects shuttle valve movement from the main oil pump supply position to the fan-driven pump oil supply position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
ENG OIL LO TEMP caution message if throttles are advanced 50%.
> 152°C (306°F)
If the oil temperature exceeds 152°C, a L (R) ENG EXCEEDENCE caution message is displayed. The ODM receives an excitation signal from the PHMU. The ODM detects ferrous and non-ferrous debris in the oil system, and sends a signal back to the PHMU. The PHMU supplies the ODM information to both channels of the EEC. If the amount of debris exceeds a predefined limit, an OIL DEBRIS ABOVE LIMIT INFO message is displayed.
< 49°C (120°F)
TECHNICAL TRAINING MANUAL
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Operation at maximum oil pressure cannot exceed 10 minutes.
ENG OIL LO PRESS warning message when oil is below minimum threshold with the engine running. CS1_CS3_TB79_002
79-38
CS130_79_L3_Indicating_DetailedDescription.fm
79 - Oil 79-30 Indicating
TO OIL TEMPERATURE PROBE
91.7
91.7
21.2
21.2
ELECTRONIC ENGINE CONTROL
N1
618 OIL PRESSURE SENSOR
OIL FILTER DIFFERENTIAL PRESSURE SENSOR
Channel A
624
ENG OIL LO TEMP EMER LTS OFF EMER DEPRESS ON HYD PUMP 2B ON APU ON CABIN PRESS MAN PARK BRAKE ON
EGT
68.2 280 72 85
N2 FF (PPH) OIL TEMP OIL PRESS
68.3 285 48 98
GEAR
DN DN DN
Channel B
IN
TOTAL FUEL (KG) 2490 SYN
OIL QUANTITY SENSOR
0
5070
SLAT / FLAP
0
2580 EICAS
OMS Excitation PROGNOSTICS AND HEALTH MANAGEMENT UNIT
HMU
OIL DEBRIS MONITOR AUXILIARY OIL PRESSURE SENSOR CS1_CS3_7930_017
LEGEND ARINC 429 CAN BUS Analog
Figure 19: Oil Indicating Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-39
CS130_79_L3_Indicating_DetailedDescription.fm
79 - Oil 79-30 Indicating
OIL PRESSURE WARNING LIMITS The EEC sends a signal to EICAS for engine oil pressure. A red L (R) ENG OIL LO PRESS EICAS warning message is shown if the oil pressure drops below the minimum threshold while the engine is running. During an oil pressure test, the oil pressure should fall within the trim band range. If the oil pressure falls between the trim band and the high oil pressure or low oil pressure yellow limit, the oil metering plugs on the oil control module (OCM) can be replaced with different sized metering plugs to bring the oil pressure into the trim band range. This can only be done after consultation with Pratt & Whitney.
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CS130_79_L3_Indicating_DetailedDescription.fm
79 - Oil 79-30 Indicating
MAIN OIL PRESSURE - MOP (PSIG)
L ENG EXCEEDANCE 200
n
eratio
al Op
orm um N
Maxim 150
tion
al Opera
m Norm Minimu 100
m Limit
Minimu 50
L ENG OIL PRESS 0 13000 (53)
14000 (57)
15000 (61)
16000 (65)
17000 (69)
18000 (74)
19000 (78)
20000 (82)
21000 (86)
22000 (90)
23000 (94)
24000 (98)
CS1_CS3_7930_020
HIGH ROTOR SPEED – N2 (RPM/1000) (% RPM)
Figure 20: Oil Pressure Warning Limits Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-41
CS130_79_L3_Indicating_DetailedDescription.fm
79 - Oil 79-30 Indicating
MONITORING AND TESTS The following page provides crew alerting system (CAS) and INFO messages for the oil indicating system.
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CS130_79_L2_Indicating_MonitoringTests.fm
79 - Oil 79-30 Indicating
CAS MESSAGES Table 2: WARNING Messages MESSAGE
LOGIC
L ENG OIL PRESS
Left engine oil pressure is below or above the normal range.
R ENG OIL PRESS
Right engine oil pressure is below or above the normal range.
Table 3: CAUTION Messages MESSAGE
LOGIC
ENG OIL LO TEMP
Either engine oil too cold to allow a high thrust to be set.
L ENG OIL FILTER
Left engine oil filter is bypassed (clogged filter).
R ENG OIL FILTER
Right engine oil filter is bypassed (clogged filter).
L ENG EXCEEDANCE
L N1 or L N2 or L EGT or L oil temperature is above the threshold.
L ENG EXCEEDANCE
R N1 or R N2 or R EGT or R oil temperature is above the threshold.
Table 4: ADVISORY Messages MESSAGE
LOGIC
L ENG OIL LO QTY
Left engine oil, low quantity detected.
R ENG OIL LO QTY
Right engine oil, low quantity detected.
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79 - Oil 79-30 Indicating
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79 - Oil 79-30 Indicating
Table 5: INFO Messages Table 5: INFO Messages MESSAGE 79 L ENGINE FAULT - OIL DEBRIS ABOVE LIMIT
MESSAGE
LOGIC INFO message will be set when an oil debris level above limit has been detected in either channel A or B by the left engine FADEC system.
79 R ENGINE FAULT - OIL DEBRIS ABOVE LIMIT
INFO message will be set when an oil debris level above limit has been detected in either channel A or B by the right engine FADEC system.
79 L ENGINE FAULT- OIL DEBRIS MON INOP
FADEC logic has determined that the sensor has failed.
79 R ENGINE FAULT- OIL DEBRIS MON INOP
FADEC logic has determined that the sensor has failed.
79 L ENGINE FAULT - OIL QTY FADEC logic has determined that the oil SNSR INOP quantity signal from channel A and/or channel B has failed.
79 L ENGINE FAULT - VORV OPER DEGRADED
FADEC logic has determined that the oil filter P signal in channel A and/or channel B has failed.
79 R ENGINE FAULT - OIL FILTER SENSOR INOP
FADEC logic has determined that the oil filter P signal in channel A and/or channel B has failed.
79 L ENGINE FAULT - OIL FILTER IMPENDING BYPASS
INFO message will be set by the left engine when the oil filter impending bypass is detected but there is no actual bypass.
79 R ENGINE FAULT - OIL FILTER IMPENDING BYPASS
INFO message will be set by the right engine when the oil filter impending bypass is detected but there is no actual bypass.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
INFO message will be set by the left engine in any of the following conditions: - Any channel of the VORV feedback sensor is failed - The active EEC channel has detected a cross-check failure on the VORV feedback sensor
79 R ENGINE FAULT - VORV OPER DEGRADED
INFO message will be set by the right engine in any of the following conditions: - Any channel of the VORV feedback sensor is failed - The active EEC channel has detected a cross-check failure on the VORV feedback sensor
79 L ENGINE FAULT - BRG DAMPER VLV INOP
INFO message will be set by the left engine in any of the following conditions: - Both EEC channels are unable to command bearing damper valve solenoid
79 R ENGINE FAULT - OIL QTY FADEC logic has determined that the oil SNSR INOP quantity signal from channel A and/or channel B has failed. 79 L ENGINE FAULT - OIL FILTER SENSOR INOP
LOGIC
- Any channel control processor failed and remaining channel unable to command bearing damper valve solenoid 79 R ENGINE FAULT - BRG DAMPER VLV INOP
TECHNICAL TRAINING MANUAL
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INFO message will be set by the right engine in any of the following conditions: - Both EEC channels are unable to command bearing damper valve solenoid - Any channel control processor failed and remaining channel unable to command bearing damper valve solenoid
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79 - Oil 79-30 Indicating
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79 - Oil 79-30 Indicating
Table 5: INFO Messages MESSAGE
LOGIC
79 L ENGINE FAULT - AUX OIL An INFO message is set by the left engine PRESS MON INOP when the auxiliary oil pressure sensor fails. This INFO message is masked if the PHMU fails. 79 R ENGINE FAULT - AUX OIL An INFO message is set by the right engine PRESS MON INOP when the auxiliary oil pressure sensor fails. This INFO message is masked if the PHMU fails.
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79 - Oil 79-30 Indicating
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CS130_79_ILB.fm
TECHNICAL TRAINING MANUAL MODULE 5: POWER PLANT CMUI: CS130-21.07-06-00-121418 VERSION V6.00 BD-500-1A10 BD-500-1A11
CUSTOMER TRAINING Montreal Training Centre 8575 Côte-de-Liesse Road Saint-Laurent, Québec, Canada H4T 1G5 Telephone (514) 344-6620 Toll-Free North America 1 (877) 551-1550 Fax (514) 344-6643 www.batraining.com iflybombardier.com
Reader Notice and Disclaimer Please note that this version of C Series Module 5: Power Plant is up-to-date as of the date of the training course(s) that you are attending, and may only be used for the purpose of such course(s). Bombardier disclaims responsibility in case of use of the document for any other purpose than said course(s). If you would like to continue using C Series Module 5: Power Plant after your subscription access has expired, you are required to register and purchase a subscription plan in order to receive the necessary updates when they become available at the following website: http://www.batraining.com. Bombardier Inc., or its subsidiaries (collectively “Bombardier”), provides this information to its customers and to government authorities in confidence. The information contained herein must therefore be treated as proprietary confidential information, and as such it must be excluded from any request for access to a record pursuant to section 20 of the Access to Information Act, RSC 1985, c A-1, or any other applicable statute with respect to access to information. Public release of this information would be highly detrimental to Bombardier and as such is strictly prohibited without Bombardier's prior written authorization. This document, which comprises protected intellectual property and trade secrets, shall not be used, reproduced, published, broadcasted, copied, translated, distributed, transferred, stored on any medium, including in a retrieval system, communicated, altered, or converted in any form or by any means, electronic or otherwise, in whole or in part, without Bombardier's prior written authorization. The rights to all patents, inventions, know-how, copyrights, trademarks, trade secrets, registered designs, database rights, semiconductor topography rights, service marks, logos, domain names, business names, trade names, moral rights and all related registrations or applications in any country or jurisdiction contained herein belong to or are used under license by Bombardier. This documentation, the technical data it contains, and all other information shall not be modified, translated, reverse assembled, reverse engineered, or decompiled and shall be used solely for training purposes. Nothing contained herein shall be constructed as granting, explicitly or implicitly, any license or other right to use the information other than for the above-stated training purposes. Copyright © 1999-2018 Bombardier Inc. or its subsidiaries. All rights reserved. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
Acronyms and Abbreviations
ACRONYMS AND ABBREVIATIONS A
AEMF
aft engine mount fitting
AFCS
automatic flight control system
AFCU
alternate flight control unit
A/T
autothrottle
AFDA
adaptive flight display application
ABIT
automatic BIT
AFDT
adaptive flight display table
ABS
absolute
AFDX
avionics full duplex switched Ethernet
ABS
autobrake system
AFM
Airplane Flight Manual
ACARS
aircraft communications addressing and reporting system
AFP
automated fiber placement
ACC
active clearance control
AGB
accessory gearbox
ACES
avionics cooling and extraction system
AGL
above ground level
ACL
access control list
AHC
attitude heading computer
ACM
air cycle machine
AHMS
aircraft health management system
ACM
aircraft condition monitoring
AIS
aircraft information server
ACMF
aircraft condition monitoring function
AIS
audio integrating system
ACMP
AC motor pump
AIM
axle interface module
ACP
audio control panel
AIM
aircraft identification module
ACU
audio conditioning unit
AIM
align-in-motion
ADC
air data computer
AL
autoland
ADEF
aircraft data exchange function
ALC
APU line contactor
ADF
automatic direction finder
ALI
airworthiness limitation item
ADI
attitude direction indicator
AI-Li
aluminum-lithium
ADLS
airborne data link system
ALM
application license manager
ADMF
aircraft data management function
ALT
altitude
ADRF
aircraft data recording function
ALTN FLAP
alternate flap
ADS
air data system
AM
amplitude modulation
ADSP
air data smart probe
AMCU
advanced monitor control unit
ADSP
air data system probe
AMP
Aircraft Maintenance Publication
AES
alternate extension system
ANR
archive noise reduction
AEV
avionics exhaust valve
ANS
aircraft network switch
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
i
Acronyms and Abbreviations
AOA
angle-of-attack
BAHX
buffer air heat exchanger
AOC
air/oil cooler
BAPS
buffer air pressure sensor
AOC
airport operational communications
BAVS
buffer air valve solenoid
AODV
active oil damper valve
BAVSOV
buffer air shutoff valve
AOHX
air/oil heat exchanger
BCCC
base coat clear coat
AP
autopilot
BCS
brake control system
APM
aircraft personality module
BDCU
brake data concentrator unit
APR
automatic power reserve
BFO
beat frequency oscillator
APU
auxiliary power unit
BGM
boarding music
AR
automatic realignment
BIT
built-in test
ARTCC
air route traffic control center
BITE
built-in test equipment
ASA
autoland status annunciator
BL
buttock line
ASC
APU starting contactor
BLC
battery line contactor
ASM
air separation module
BLS
bifurcation latch system
ASRP
aircraft structure repair publication
BMPS
bleed monitoring pressure sensor
AT
autothrottle
BPCU
bus power control unit
ATC
air traffic control
BPMS
bleed pressure monitoring sensor
ATIS
air traffic information services
BSC
battery start contactor
ATP
acceptance test procedure
BTC
bus tie contactor
ATS
air turbine starter
BTS
base transceiver station
ATS
autothrottle system
BTMS
brake temperature monitoring system
ATS
air traffic services
BTS
brake temperature sensor
AV-VENTS
avionics ventilated temperature sensor
BTS
bleed temperature sensor
BVID
barely visible impact damage
B BALODS
bleed air leak and overheat detection system
BAP
buffer air pressure
CADTS
cargo duct temperature sensor
BAS
bleed air system
CAI
cowl anti-ice
BAV
bleed air valve
CAIS
cowl anti-ice system
BACV
buffer air check valve
CAIV
cowl anti-ice valve
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
C
TECHNICAL TRAINING MANUAL
For Training Purposes Only
ii
Acronyms and Abbreviations
CAM
cockpit area microphone
CMUI
configuration management unique identifier
CAN
controller air network
CNS
communications, navigation and surveillance
CAS
calibrated airspeed
COM
communication
CAS
crew alerting system
CPCS
cabin pressure control system
CATS
cargo temperature sensor
CPCV
compensator pressure check valve
CB
circuit breaker
CPD
circuit protection device
CBIT
continuous built-in test
CPDD
circuit protection device detector
CBP
circuit breaker panel
CPDLC
controller-pilot data link communications
CBV
cross-bleed valve
CPN
Collins part number
CC
cabin controller
CPU
central processing unit
CCDL
cross-channel data link
CRC
cyclical redundancy check
CCM
common computing module
CRES
corrosion resistant steel
CCMR
common computing module runtime
CSD
cabin service display
CCP
cursor control panel
CSD
customer service display
CCU
camera control unit
CSMU
crash-survivable memory unit
CCW
counterclockwise
CSOV
cargo shutoff valve
CDC
control and distribution cabinet
CT
crew terminal
CDI
course deviation indicator
CT
current transformer
CDTS
compressor discharge temperature sensor
CTP
control tuning panel
CEM
cover and environmental module
CVR
cockpit voice recorder
CF
configuration file
CW
clockwise
CFIT
controlled flight into terrain
CWB
center wing box
CFRP
carbon fiber reinforced polymer
CIC
compressor intermediate case
D/I
discrete input
CIC
corrosion inhibiting compound
D/O
discrete output
CLAWS
control laws
DBM
database manager
CM
configuration manager
DCM
data concentrator module
CMS
cabin management system
DCMR
data concentration module runtime
CMU
communication management unit
DCS
data concentration system
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
D
TECHNICAL TRAINING MANUAL
For Training Purposes Only
iii
Acronyms and Abbreviations
DCU
directional control unit
EFAN
extraction fan
DCV
directional control valve
EDM
emergency descent mode
DFSOV
dual-flow shutoff valve
EDP
engine-driven hydraulic pump
DLCA
data link communications application
EDP
engine-driven pump
DMA
display manager application
EDU
electronic display unit
DMC
data concentrator unit module cabinet
EEC
electronic engine control
DME
distance measuring equipment
EEGS
emergency electrical power generation
DMM
data memory module
EEPROM
electrical erasable programmable read only memory
DP
differential protection
EESS
emergency escape slide system
DPCT
differential protection current transformer
EFB
electronic flight bag
DPI
differential pressure indicator
EEGS
emergency electrical power generation
DPLY
deploy
EFCS
electronic flight control system
DRA
diagnostic and reporting application
eFIM
electronic fault isolation manual
DSK
double-stack knob
EFIS
electronic flight instrument system
DSM
digital switching module
EGT
exhaust gas temperature
DSU
data storage unit
EHSV
electrohydraulic servovalve
DSPU
diode shunt protection unit
EIC
engine inlet cowl
DTC
DC tie contactor
EICAS
engine indication and crew alerting system
DTE
damage tolerance evaluation
ELC
external power line contactor
DTI
damage tolerance inspection
ELT
emergency locator transmitter
DTS
duct temperature sensor
EMA
electric motor actuator
DU
display unit
EMU
expansion module unit
E
EMAC
electric motor actuator controller
EBC
essential bus contactor
EMCU
electric motor control unit
ECL
electronic checklist
EMER
emergency
ECS
environmental control system
EMPC
emergency power control
ECU
electronic control unit
EOAM
emergency opening assist means
ECU
external compensation unit
EOF
end-of-flight
EDCM
electronic door control module
EPC
electrical power center
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
iv
Acronyms and Abbreviations
EPC
electronic power center
FCV
flow control valve
EPGD
electrical power generation and distribution
FD
flight director
EPGDS
electrical power generation and distribution system
FDDSS
flight deck door surveillance system
EPGS
electrical power generation system
FDE
flight deck effect
EPP
engine programming plug
FDG
fan drive gearbox
EPSU
emergency power supply unit
FDGS
fan drive gear system
EPTRU
external power transformer rectifier unit
FDR
flight data recorder
ERAV
emergency ram air valve
FDRAS
flight deck remote access system
ESD
electrostatic discharge
FDV
flow divider valve
ETC
essential tie contactor
FEGV
fan exit guide vane
ETOPS
extended-range twin-engine operational performance standards
FEMB
forward engine mount bulkhead
FF
fuel flow
F
FFDP
fuel flow differential pressure
FA
flight attendant
FFSV
free fall selector valve
FAA
Federal Aviation Authority
FG
flight guidance
FADEC
full authority digital engine control
FGS
flight guidance system
FANS
future air navigation system
FIC
fan intermediate case
FAV
fan air valve
FIDEX
fire detection and extinguishing
FBC
front bearing compartment
FIM
fault isolation manual
FBW
fly-by-wire
FLC
flight level change
FBWPC
fly-by-wire power converter
FLS
fast load-shed
FC
full close
FLTA
forward-looking terrain avoidance
FCP
flight control panel
FMA
flight mode annunciator
FCS
flight control system
FMS
flight management system
FCBS
fatigue critical baseline structure
FMSA
flight management system application
FCEE
flight crew emergency exit
FMV
fuel metering valve
FCSB
fan cowling support beam
FO
full open
FCU
flush control unit
FOD
foreign object debris
FCU
fuel control unit
FOHX
fuel/oil heat exchanger
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
v
Acronyms and Abbreviations
FOHXBV
fuel/oil heat exchanger bypass valve
GSA
ground spoiler actuator
FPS
feedback position sensor
GSCM
ground spoiler control module
FPV
flight path vector
GSE
ground support equipment
FQC
fuel quality computer
GUI
graphical user interface
FS
fixed structure
FS
flight station
HAAO
high altitude airport operation
FS
fuselage station
HDG
heading
FSA
file server application
HF
high frequency
FSV
flow sensor venturi
HID
high-intensity discharge
FSB
fasten seat belt
HLEIF
high load event indication function
FSCL
flight spoiler control lever
HLSL
high lift selector lever
FTIS
fuel tank inerting system
HMU
health management unit
FW
failure warning
HOR
hold open rod
FWSOV
firewall shutoff valve
HP
high-pressure
HPC
high-pressure compressor
H
G GA
go-around
HPD
hydraulic pump depressurization
GCF
ground cooling fan
HPGC
high-pressure ground connection
GCR
generator control relay
HPSOV
high-pressure shutoff valve
GCS
global connectivity suite
HPT
high-pressure turbine
GCU
generator control unit
HPV
high-pressure valve
GFP
graphical flight planning
HRD
high-rate discharge
GFRP
glass fiber reinforced polymer
HRTDb
high-resolution terrain database
GHTS
galley heater temperature sensor
HS
handset
GLC
generator line contactor
HS
high solid
GMT
Greenwich mean time
HSI
horizontal situation indicator
GNSS
global navigation satellite system
HSTA
horizontal stabilizer trim actuator
GPWS
ground proximity warning system
HSTS
horizontal stabilizer trim system
GS
glideslope
HUD
head-up display
GS
ground spoiler
HUDS
head-up display system
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
vi
Acronyms and Abbreviations
I
IOM
input/output module
I/O
input/output
IP
information provider
IAMS
integrated air management system
IPC
integrated processing cabinet
IAS
indicated airspeed
IPCKV
intermediate pressure check valve
IASC
integrated air system controller
IPS
inches per second
IBIT
initiated built-in test
IPS
integrated processing system
IBR
integral bladed rotor
IRCV
inlet return check valve
ICAO
International Civil Aviation Organization
IRS
inertial reference system
ICCP
integrated cockpit control panel
IRU
inertial reference unit
ICDU
integrated control display unit
ISI
integrated standby instrument
ICU
inerting control unit
ISM
input signal management
ICU
isolation control unit
ISPS
in-seat power supply
ICV
isolation control valve
ISPSS
in-seat power supply system
IFE
in-flight entertainment system
ITT
interturbine temperature
IFEC
in-flight entertainment and connectivity system
IFIS
integrated flight information system
IFPC
integrated fuel pump and control
IFS
information landing system
L/S
lube/scavenge
IFS
inner fixed structure
LAN
local area network
IGN
ignition
LBIT
landing built-in test
IGV
inlet guide vane
LCD
life cycle data
IGVA
inlet guide vane actuator
LCT
line current transformer
IIM
inceptor interface module
LED
light-emitting diode
IIV
inlet isolation valve
LLU
LED lighting unit
ILS
instrument landing system
LGCL
landing gear control lever
IMA
integrated modular avionics
LGCV
landing gear control valve
IMS
information management system
LGIS
landing gear indicating system
INT
intermittent
LGSCU
landing gear and steering control unit
IOC
input/output concentrator
LGSV
landing gear selector valve
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
J JOSV
journal oil shuttle valve L
TECHNICAL TRAINING MANUAL
For Training Purposes Only
vii
Acronyms and Abbreviations
LOC
localizer
MIXTS
mix manifold temperature sensor
LOP
low oil pressure
MKP
multifunction keyboard panel
LOPA
layout of passenger area
MLG
main landing gear
LP
low-pressure
MLW
maximum landing weight
LPC
low-pressure compressor
MMEL
Master Minimum Equipment List
LPSOV
low-pressure shutoff valve
MOF
main oil filter
LPT
low-pressure turbine
MOT
main oil temperature
LRD
low-rate discharge
MPP
maintenance planning publication
LRM
line replaceable module
MPSOV
minimum pressure-shutoff valve
LRU
line replaceable unit
MRW
maximum ramp weight
LSK
line select key
MSV
mode select valve
LSOP
lubrication and scavenge oil pump
MTD
master time and date
LV
lower sideband voice
MTO
maximum rated takeoff
LVDS
low-voltage differential signaling
MTOW
maximum takeoff weight
LVDT
linear variable differential transformer
MWW
main wheel well
MZFW
maximum zero fuel weight
M MAX
maximum
MB
marker beacon
NA
not activated
MCDL
motor control data link
NACA
National Advisory Committee for Aeronautics
MCE
motor control electronic
ND
nosedown
MCR
minimum control requirement
NCD
no computed data
MCV
mode control valve
NCG
network communication gap
MDU
manual drive unit
NCU
network control unit
MEL
minimum equipment list
NDB
non-directional beacon
MES
main engine start
NDO
network data object
MFK
multifunction keyboard panel
NEA
nitrogen-enriched air
MFP
multifunction probe
NEADS
nitrogen-enriched air distribution system
MFS
multifunction spoiler
NLG
nose landing gear
MFW
multifunction window
NO PED
no personal electronic device
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
N
TECHNICAL TRAINING MANUAL
For Training Purposes Only
viii
Acronyms and Abbreviations
NPRV
negative pressure-relief valve
NU
noseup
P&W
Pratt and Whitney
NWS
nosewheel steering
P2
inlet pressure
NVM
non-volatile memory
PA
passenger address
O
PAX
passenger
OAT
outside air temperature
PBA
pushbutton annunciator
OBB
outboard brake
PBE
protective breathing equipment
OBIGGS
onboard inlet gas generation system
PBIT
power-up built-in test
OC
overcurrent
PCE
precooler exhaust
OCM
oil control module
PCE
precooler exit
OCM
option control module
PCU
power control unit
ODI
overboard discharge indicator
portable document format
ODL
onboard data loader
PDOS
power door operating system
ODM
oil debris monitor
PDPS
pack pressure differential sensor
OEA
oxygen-enriched air
PDL
permitted damage limits
OEM
original equipment manufacturer
PDS
power distribution system
OF
overfrequency
PDTS
pack discharge temperature sensor
OFV
outflow valve
PDU
power drive unit
OMS
onboard maintenance system
PED
personal electronic device
OMS IMA
OMS interactive maintenance application
PEM
power environment module
OMSA
onboard maintenance system application
PEV
pressure equalization valve
OMST
onboard maintenance system table
PFCC
primary flight control computer
OPAS
outboard position asymmetry sensor
PFD
primary flight display
OPU
overvoltage protection unit
PFS
post flight summary
OSP
opposite-side pressure
PHMU
prognostics and health management unit
OSS
overspeed/shutdown solenoid
PIC
peripheral interface controller
OT
other traffic
PIC
processor-in-command
OV
overvoltage
PIFS
pack inlet flow sensor
OWEE
overwing emergency exit
PIM
panel interface module
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
P
TECHNICAL TRAINING MANUAL
For Training Purposes Only
ix
Acronyms and Abbreviations
PIPS
pack inlet pressure sensor
PVT
position, velocity, time
PLD
programmable logic device
PWM
pulse width modulation
PLD
proportional lift dump
PMA
permanent magnet alternator
QAD
quick attach/detach
PMA
program manager application
QEC
quick engine change
PMAG
permanent magnet alternator generator
PMG
permanent magnet generator
RA
radio altimeter
POB
power off brake
RA
resolution advisory
POB
pressure off brake
RAM
receiver autonomous integrity monitoring
POR
point of regulation
RAD
radio altitude
PPM
power producing module
RARV
ram air regulating valve
PPT
pedal position transducer
RAT
ram air turbine
PRAM
prerecorded announcement and message
RDC
remote data concentrator
PRSOV
pressure-regulating shutoff valve
RDCP
refuel/defuel control panel
PRV
pressure-regulating valve
REL
relative altitude
PRV
pressure-relief valve
REO
repair engineering order
PS
passenger service
RET
retracted
PS
pressure sensor
REU
remote electronic unit
PSA
print server application
RF
radio frequency
PSE
principal structural element
RFAN
recirculation fan
PSU
passenger service unit
RGA
rotary geared actuator
PSUC
passenger service unit controller
RGC
RAT generator control
PT
proximate traffic
RIPS
recorder independent power supply
PT
pressure transducer
RIU
radio interface unit
Pt
total pressure
RLC
RAT line contactor
PTS
pack temperature sensor
RMA
remote maintenance access
PTT
push-to-talk
RMS
radio management system
PTU
power transfer unit
ROLS
remote oil lever sensor
PTY
priority
ROV
redundant overvoltage
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Q
R
TECHNICAL TRAINING MANUAL
For Training Purposes Only
x
Acronyms and Abbreviations
RPA
rudder pedal assembly
SFV
safety valve
RPM
revolutions per minute
SLS
slow load-shed
RSA
report server application
SMS
surface management system
RSP
reversion switch panel
SOV
solenoid operated valve
RTA
receiver-transmitter antenna
SOV
shutoff valve
RTD
resistance temperature device
SPCV
supply pressure check valve
RTD
resistive thermal device
SPDS
secondary power distribution system
RTL
ready-to-load
SPDT
single pole double throw
RTO
rejected takeoff
SPKR
speaker
RTS
return to service
SPM
seat power module
RTSA
radio tuning system application
SSC
sidestick controller
RVDT
rotary variable differential transformer
SSD OML
solid-state onboard media loader
SSI
structural significant item
S SAL
specific airworthiness limitation
SSEC
static source error connection
SAT
static air temperature
SSPC
solid-state power controller
SATCOM
satellite communication
SSPC-CB
solid-state power controller circuit breaker
SAV
starter air valve
SSRPC
solid-state remote power controller
SB
service bulletin
SUA
special use airspace
SBAS
satellite-based augmentation system
SVA
stator vane actuator
SBIT
start-up BIT
SVS
synthetic vision system
SCV
surge control valve
SCV
steering control valve
T/M
torque motor
SEB
seat electronics box
T/R
thrust reverser
SELCAL
selective calling
T2
inlet temperature
SFCC
slat/flap control computer
TA
traffic advisory
SFCL
slat/flap control lever
TACKV
trim air check valve
SFCP
slat/flap control panel
TAPRV
trim air pressure-regulating valve
SFECU
slat/flap electronic control unit
TASOV
trim air shutoff valve
SFIS
standby flight instrument system
TAT
total air temperature
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
T
TECHNICAL TRAINING MANUAL
For Training Purposes Only
xi
Acronyms and Abbreviations
TAV
trim air valve
TQA
throttle quadrant assembly
TAWS
terrain awareness and warning system
TRAS
thrust reverser actuation system
TAWSDb
terrain awareness and warning system database
TRU
transformer rectifier unit
TCA
turbine cooling air
TSC
TRU start contactor
TCAS
traffic alert and collision avoidance system
TSFC
thrust specific fuel consumption
TCB
thermal circuit breaker
TSM
trip status monitor
TCDS
type certificate data sheet
TSO
technical standard order
TCF
terrain control valve
TSS
traffic surveillance system
TDR
transponder
TTG
time-to-go
TCV
temperature control valve
TTP
time-triggered protocol
TEC
turbine exhaust case
TWIP
terminal weather information for pilot
TED
trailing edge down
TEU
trailing edge up
UART
universal asynchronous receiver transmitter
TFTP
trivial file transfer protocol
UBMF
usage-based monitoring function
TIC
turbine inlet case
UF
underfrequency
TIC
turbine intermediate case
ULB
underwater locator beacon
TIV
temperature inlet valve
UPLS
ultrasonic point level sensors
TIV
temperature isolation valve
USB
universal serial bus
TLA
throttle lever angle
UTC
universal time coordinated
TLC
TRU line contactor
UV
upper sideband voice
TLD
time limited dispatch
UV
ultraviolet
TOGA
takeoff/go-around
UV
undervoltage
TPIS
tire pressure indicating system
TPM
TAWS processing module
VAC
volts alternating current
TPM
tire pressure module
VDC
voltage direct current
TPMA
terrain processing module application
VDL
VHF data link
TPMU
tire pressure monitoring unit
VDLM
VHF data link mode
TPS
tire pressure sensor
VENTS
ventilated temperature sensor
TPSA
terrain processing system application
VFG
variable frequency generator
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
U
V
TECHNICAL TRAINING MANUAL
For Training Purposes Only
xii
Acronyms and Abbreviations
VFGOOHX
variable frequency generator oil/oil heat exchanger
WST
wheel speed transducer
VGMD
vacuum generator motor drive unit
WTBF
wing-to-body fairing
VHF-NAV
VHF navigation
WWHS
windshield and side window heating system
VID
visible impact damage
WWS
waste water system
VL
virtual link
WWSC
water and waste system controller
VLAN
virtual local area network
WXR
weather radar
VNAV
vertical navigation
VOC
volatile organic compounds
ZB
zone box
VOR-VHF
VHF omnidirectional radio
P
differential pressure
VORV
variable oil reduction valve
VPA
video passenger announcement
VSD
vertical situation display
VSPD
V-speed
VSWR
voltage standing-wave radio
VTU
video transmission unit
Z
W WAI
wing anti-ice
WAP
wireless access point
WAIS
wing anti-ice system
WAITS
wing anti-ice temperature sensor
WAIV
wing anti-ice valve
WBV
windmill bypass valve
WIPC
windshield ice protection controller
WL
waterline
WOFFW
weight-off-wheels
WOW
weight-on-wheels
WPS
words-per-second
WS
wing situation
WSA
web server application
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
xiii
Acronyms and Abbreviations
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
xiv
Module 5: Power Plant - List of Changes
MODULE 5: POWER PLANT - LIST OF CHANGES The following table details the changes applied to this revision: ATA NAME AND NUMBER
CMUI NUMBER
CHANGES APPLIED
Front Page
CS130-21.07-06-00-121418
Version number and CMUI number
Acronyms and Abbreviations
CS130-21.07-06-00-121418
Deleted BCT (bus tie contactor current transformer) and added MKP (multifunction keyboard panel) (pges ii and viii)
List of Changes
CS130-21.07-06-00-121418
New CMUI
21 Environmental Control
CS130-21.07-06-00-121418
No change
28 Fuel
CS130-21.07-06-00-121418
Added new Fuel Safety section (pges 28-2 to 28-13) Updated collector tank area (pge 28-21) Updated graphic (pges 28-31, 28-33, 28-35) Added maximum suction pressure (pge 28-36) Updated graphic (pges 28-37, 28-39, 28-71, 28-87, 28-101, 28-103, 28-105) Added caution (pge 28-106) Updated graphic (pges 28-107, 28-109, 28-111)
30 Ice and Rain Protection
CS130-21.07-06-00-121418
Moved System Test before CAS Messages (pges 30-30 to 30-35)
36 Pneumatic
CS130-21.07-06-00-121418
No change
47 Liquid Nitrogen (Fuel Tank Inerting System)
CS130-21.07-06-00-121418
No change
49 Airborne Auxiliary Power
CS130-21.07-06-00-121418
Updated graphic (pges 49-9, 49-13) Added note and caution (pge 49-10) Updated text (pge 49-14) Updated graphic (pges 49-15, 49-23, 49-97)
71 Power Plant
CS130-21.07-06-00-121418
No change
72 Engine
CS130-21.07-06-00-121418
Updated graphic (pge 72-3)
73 Engine Fuel and Control
CS130-21.07-06-00-121418
Changed throttle lever angle to thrust lever angle (pge 73-43)
74-80 Ignition and Starting
CS130-21.07-06-00-121418
No change
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
LOC i
Module 5: Power Plant - List of Changes
ATA NAME AND NUMBER
CMUI NUMBER
CHANGES APPLIED
75 Engine Air Systems
CS130-21.07-06-00-121418
No change
76 Engine Controls
CS130-21.07-06-00-121418
No change
77 Engine Indicating
CS130-21.07-06-00-121418
Updated graphic; added indicator for fan drive shaft dimple location, and view forward looking aft (pge 77-43) Added new Fan Trim Balance Using the PW 1500G Trim Balance Helper section with graphics (pges 77-58 to 77-69)
78 Exhaust
CS130-21.07-06-00-121418
Updated graphic (pge 78-41) Updated text (pge 78-42) Changed ICV control solenoid to ICU isolation valve control solenoid (pge 78-44) Changed valve to unit in figure title (pge 78-45) Changed DCV energized to DCU energized in heading and figure title (pges 78-48, 78-49) Changed DCV control solenoid to DCU control solenoid (pge 78-52) Changed ICV control solenoid to ICU control solenoid (pge 78-54) Changed ICV and DCV enable solenoids to ICU and DCU control solenoids; ICV solenoid to ICU solenoid; DCV solenoid to DCU solenoid; added information for maximum reverse thrust (pge 78-56)
79 Oil
CS130-21.07-06-00-121418
Added text for oil level and updated caution (pge 79-4) Updated graphic (pge 79-23) Updated tables in graphic (pge 79-37) Changed 47°C (117°F) to 49°C (120°F) (pge 79-38) Updated EICAS on graphic (pge 79-39) Updated graphic (pge 79-41)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
LOC ii
ATA 21 - Environmental Control
BD-500-1A10 BD-500-1A11
21 - Environmental Control
Table of Contents 21-90 Integrated Air System Controller ..................................21-2
21-51 Cooling....................................................................... 21-34
General Description ..........................................................21-2 Component Location .......................................................21-4 Detailed Description .........................................................21-6 Monitoring and Tests .......................................................21-8 INFO Messages ...........................................................21-9 Practical Aspects.............................................................21-10 Onboard Maintenance System Test ..........................21-10
General Description ...................................................... 21-34 Component Location .................................................... 21-36 Water Extractor ......................................................... 21-38 Water Sprayer ........................................................... 21-38 Pack Discharge Pressure Sensor ............................. 21-40 Component Information ................................................ 21-42 Temperature Control Valve ....................................... 21-42 Ram Air Regulating Valve ........................................ 21-44 Controls and Indications ................................................ 21-46 Controls..................................................................... 21-46 Indications ................................................................. 21-48 Detailed Description ..................................................... 21-50 Pack Temperature Regulation .................................. 21-50 Pack Icing Protection ............................................... 21-50 Pack Overheat Protection ........................................ 21-50 Pack Discharge Overheat Protection ....................... 21-50 Monitoring and Tests .................................................... 21-52 CAS Messages ......................................................... 21-53
21-53 Flow Control System ..................................................21-12 General Description .......................................................21-12 Component Location .....................................................21-14 Flow Sensor Venturi...................................................21-14 Flow Control Valve.....................................................21-14 Ozone Converter/Ozone and VOC Converter (Optional)..........................................21-14 Pack Inlet Pressure Sensors......................................21-16 Pack Inlet Flow Sensors ............................................21-16 Controls and Indications ................................................21-18 Controls......................................................................21-18 Indications..................................................................21-20 Detailed Description .......................................................21-22 Flow Control Normal Operation .................................21-22 Flow Control...............................................................21-24 Flow Control Modes ...................................................21-26 Flow Control Shutoff ..................................................21-28 Flow Control Valve Failure.........................................21-28 Monitoring and Tests ......................................................21-30 CAS Messages ..........................................................21-31 Practical Aspects.............................................................21-32
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
21-52 Ram Air System......................................................... 21-56 General Description ...................................................... 21-56 Component Location .................................................... 21-58 Emergency Ram Air Valve ........................................ 21-58 Controls and Indications ................................................ 21-60 Detailed Description ..................................................... 21-62 Monitoring and Tests .................................................... 21-64 CAS Messages ......................................................... 21-65 21-60 Trim Air System ......................................................... 21-66 General Description ...................................................... 21-66
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_M5_21_EnviroContTOC.fm
21 - Environmental Control
Component Location .....................................................21-68 Trim Air Valves...........................................................21-70 Controls and Indications ................................................21-72 Detailed Description ......................................................21-74 Trim Air Control..........................................................21-74 Monitoring and Tests ......................................................21-76 CAS Messages ..........................................................21-77 21-60 Temperature Control System .....................................21-78 General Description .......................................................21-78 Component Location ......................................................21-80 Mix Manifold Temperature Sensors ..........................21-80 Duct Temperature Sensors .......................................21-82 Ventilated Temperature Sensors ..............................21-84 Cargo Duct Temperature Sensor...............................21-86 Cargo Temperature Sensor ......................................21-86 Component Information ..................................................21-88 Ventilated Temperature Sensor .................................21-88 Controls and Indications .................................................21-90 AIR Panel...................................................................21-90 AIR Synoptic Page.....................................................21-92 Cabin Management System Temperature Controls ................................................21-94 Detailed Description ......................................................21-96 Automatic Temperature Control.................................21-96 Manual Temperature Control .....................................21-96 FWD Cargo Temperature Control..............................21-98 Monitoring and Tests ....................................................21-100 CAS Messages ........................................................21-101 21-22 Recirculation.............................................................21-102 General Description .....................................................21-102 Component Location ...................................................21-104 Component Information ...............................................21-106 Recirculation Filters .................................................21-106 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Controls and Indications ............................................. 21-108 Detailed Description ................................................... 21-110 Monitoring and Test .................................................... 21-112 CAS Messages ....................................................... 21-113 21-22 Distribution System.................................................. 21-114 General Description .................................................... 21-114 Zone Distribution ..................................................... 21-116 Overhead Bin Distribution ....................................... 21-120 Flight Deck Distribution ........................................... 21-122 Component Location .................................................. 21-124 Component Information .............................................. 21-126 Low-Pressure Ground Connection.......................... 21-126 Mix Manifold ............................................................ 21-128 21-23 Ventilation System ................................................... 21-130 General Description .................................................... 21-130 Component Location .................................................. 21-132 Controls and Indications .............................................. 21-134 AIR Panel ................................................................ 21-134 Cargo Panel ............................................................ 21-134 Indications ............................................................... 21-136 Detailed Description .................................................... 21-138 Monitoring and Tests ................................................... 21-140 CAS Messages ....................................................... 21-141 Practical Aspects ......................................................... 21-142 21-40 Heating .................................................................... 21-144 Flight Crew Floor Heating ................................................. 21-144 General Description ..................................................... 21-144 Component Location .................................................. 21-144 Heated Mat.............................................................. 21-144 Detailed Description ................................................... 21-146
TECHNICAL TRAINING MANUAL
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CS130_M5_21_EnviroContTOC.fm
21 - Environmental Control
Galley Heating ...................................................................21-148 General Description ......................................................21-148 Component Location ...................................................21-150 Galley Fan................................................................21-150 Galley Heater ...........................................................21-150 Galley Heater Temperature Sensor .........................21-150 Controls and Indications ..............................................21-152 Galley Heat Controls................................................21-152 Galley Heat Indications ...........................................21-154 Detailed Description ....................................................21-156 21-26 Avionics Cooling and Extraction System ..................21-158 General Description .....................................................21-158 Air Inlet Cooling System ....................................................21-160 General Description .....................................................21-160 Component Location ...................................................21-162 Forward Avionics Filter ............................................21-162 Forward Bay Fan .....................................................21-162 Forward Skin Heat Exchanger .................................21-162 Ground Valve ...........................................................21-162 Mid Bay Fan.............................................................21-162 Avionics Coalescent Filter........................................21-162 Mid Skin Heat Exchanger ........................................21-162 Mid Avionics Filter....................................................21-162 Ducting.....................................................................21-162 Component Information ................................................21-164 Forward Avionics Filter ............................................21-164 Forward Skin Heat Exchanger .................................21-164 Ground Valve ...........................................................21-166 Avionics Coalescent Filter........................................21-166 Mid Skin Heat Exchanger ........................................21-168 Mid Avionics Filter....................................................21-168 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Controls and Indications ............................................. 21-170 Detailed Description ................................................... 21-172 Air Extraction System ........................................................ 21-174 General Description .................................................... 21-174 Component Location ................................................... 21-176 Extraction Fans ....................................................... 21-176 Ducting .................................................................... 21-176 Detailed Description ................................................... 21-178 Backup Exhaust and Extraction System ........................... 21-180 General Description .................................................... 21-180 Component Location .................................................. 21-182 Avionics Exhaust Valves ......................................... 21-182 Backup Fan ............................................................. 21-182 Ducting .................................................................... 21-182 Avionics Bay Ventilated Temperature Sensors ....... 21-184 Detailed Description .................................................... 21-186 Component Information .............................................. 21-188 Avionics Exhaust Valves ......................................... 21-188 Backup Fan ............................................................. 21-190 Controls and Indications ............................................. 21-192 Monitoring and Tests .................................................. 21-194 CAS Messages ....................................................... 21-195 Wing-to-Body Fairing Ventilation System ......................... 21-198 General Description ..................................................... 21-198 High-Pressure Pneumatic Duct Area Cooling ......... 21-200 21-30 Cabin Pressure Control System .............................. 21-202 General Description .................................................... 21-202 Automatic Mode ...................................................... 21-202 Manual Mode .......................................................... 21-202 Ditching ................................................................... 21-202
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_M5_21_EnviroContTOC.fm
21 - Environmental Control
Emergency Depressurization...................................21-202 Safety Functions ......................................................21-202 Typical Flight Profile.................................................21-204 Component Location ...................................................21-206 Outflow Valve...........................................................21-206 Water Activated Check Valve ..................................21-206 Muffler ......................................................................21-206 Safety Valve.............................................................21-206 Negative Pressure-Relief Valve ...............................21-206 Pressure Equalization Valves ..................................21-206 Component Information ...............................................21-208 Outflow Valve...........................................................21-208 Water Activated Check Valve .................................21-210 Muffler ......................................................................21-210 Safety Valve.............................................................21-212 Negative Pressure-Relief Valve ...............................21-212 Pressure Equalization Valves ..................................21-214 Detailed Component Information .................................21-216 Integrated Air System Controller..............................21-216 Controls and Indications ..............................................21-218 Pressurization Panel ................................................21-218 FMS Page ................................................................21-220 Indications................................................................21-222 Detailed Description ....................................................21-224 Auto/Manual Selection .............................................21-224 Interface ...................................................................21-226 Cabin Altitude Limitation Annunciation ....................21-228 Emergency Depressurization...................................21-230 Ditching ....................................................................21-232 Monitoring and Tests ...................................................21-234 CAS Messages ........................................................21-235
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_M5_21_EnviroContTOC.fm
21 - Environmental Control
List of Figures Figure 1: Integrated Air System Controller ...........................21-3 Figure 2: Integrated Air System Controller Location.............21-5 Figure 3: Integrated Air System Controller Detailed Description ..............................................21-7 Figure 4: Environmental Control System OMS Page .........21-11 Figure 5: Flow Control System ...........................................21-13 Figure 6: Flow Sensor Venturi, Flow Control Valve, and Ozone Converter Location ...........................21-15 Figure 7: PIP Sensor and PIF Sensor Locations................21-17 Figure 8: Flow Control Controls..........................................21-19 Figure 9: Flow Control Indications ......................................21-21 Figure 10: Flow Control Normal Operation...........................21-23 Figure 11: Flow Control Valve ..............................................21-25 Figure 12: Flow Control Priority ............................................21-27 Figure 13: Flow Control Shutoff............................................21-29 Figure 14: Flow Control Valve Lockout.................................21-33 Figure 15: Air Conditioning System ......................................21-35 Figure 16: Air Conditioning Components Location ...............21-37 Figure 17: Water Extractor and Water Sprayer ....................21-39 Figure 18: Pack Discharge Pressure Sensor Location........21-41 Figure 19: Temperature Control Valve .................................21-43 Figure 20: Ram Air Regulating Valve ...................................21-45 Figure 21: Air Conditioning Pack Controls............................21-47 Figure 22: Air Conditioning Pack Indications........................21-49 Figure 23: Air Conditioning System Detailed Description ............................................21-51 Figure 24: Ram Air System ..................................................21-57 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 25: Emergency Ram Air Valve.................................. 21-59 Figure 26: Ram Air System Controls and Indications .......... 21-61 Figure 27: Ram Air System Schematic ................................ 21-63 Figure 28: Trim Air System .................................................. 21-67 Figure 29: Trim Air Shutoff Valves, Trim Air Check Valves, and Trim Air Pressure-Regulating Valve Location..... 21-69 Figure 30: Trim Air Valve Locations..................................... 21-71 Figure 31: Trim Air System Controls and Indications........... 21-73 Figure 32: Trim Air Control................................................... 21-75 Figure 33: Temperature Control System.............................. 21-79 Figure 34: Mix Manifold Temperature Sensors.................... 21-81 Figure 35: Duct Temperature Sensors................................. 21-83 Figure 36: Flight Deck and Cabin Ventilated Temperature Sensors ......................................... 21-85 Figure 37: Cargo Duct Temperature Sensor and Cargo Temperature Sensor ......................... 21-87 Figure 38: Ventilated Temperature Sensor.......................... 21-89 Figure 39: Temperature System Controls............................ 21-91 Figure 40: Temperature System Indications ........................ 21-93 Figure 41: Cabin Management System Temperature Controls......................................... 21-95 Figure 42: Flight Deck, Forward Cabin, and Aft Cabin Temperature Control System.............. 21-97 Figure 43: Cargo Temperature Control................................ 21-99 Figure 44: Recirculation System ........................................ 21-103 Figure 45: Recirculation System Component Location...... 21-105
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-v
CS130_M5_21_EnviroContLOF.fm
21 - Environmental Control
Figure 46: Recirculation Filter.............................................21-107 Figure 47: Recirculation System Controls and Indications....................................21-109 Figure 48: Recirculation Fan Detailed Description .............21-111 Figure 49: Distribution System General Description...........21-115 Figure 50: Zone Distribution CS100 ...................................21-117 Figure 51: Zone Distribution CS300 ...................................21-119 Figure 52: Overhead Bin Distribution .................................21-121 Figure 53: Flight Deck Distribution .....................................21-123 Figure 54: Distribution System Component Location .........21-125 Figure 55: Low-Pressure Ground Connection ....................21-127 Figure 56: Mix Manifold ......................................................21-129 Figure 57: Ventilation System Component Location ..........21-131 Figure 58: Cargo Shutoff Valves Component Location ......21-133 Figure 59: Ventilation System Controls ..............................21-135 Figure 60: Ventilation System Controls and Indications .....21-137 Figure 61: Ventilation System Detailed Description ...........21-139 Figure 62: Cargo Shutoff Valve Practical Aspects..............21-143 Figure 63: Flight Crew Floor Heating General Description...........................................21-145 Figure 64: Flight Crew Floor Heating Detailed Description ..........................................21-147 Figure 65: Galley Heating...................................................21-149 Figure 66: Supplemental Galley Heat Component Location .........................................21-151 Figure 67: Galley Heater Controls ......................................21-153 Figure 68: Galley Heater Indications ..................................21-155 Figure 69: Supplemental Galley Heat Detailed Description ..........................................21-157 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 70: Avionics Cooling and Extraction System General Description .......................................... 21-159 Figure 71: Air Inlet Cooling Schematic............................... 21-161 Figure 72: Air Inlet Cooling Component Location .............. 21-163 Figure 73: Forward Avionics Filter and Forward Skin Heat Exchanger.......................... 21-165 Figure 74: Ground Valve and Avionics Coalescent Filter............................................... 21-167 Figure 75: Mid Skin Heat Exchanger and Mid Avionics Filter............................................. 21-169 Figure 76: Avionics Cooling and Extraction System Controls .............................. 21-171 Figure 77: Air Inlet Cooling System ................................... 21-173 Figure 78: Air Extraction System ....................................... 21-175 Figure 79: Main Air Extraction System Component Location......................................... 21-177 Figure 80: Main Air Extraction System............................... 21-179 Figure 81: Backup Exhaust and Air Extraction System General Description .......................................... 21-181 Figure 82: Backup Exhaust and Air Extraction System .... 21-183 Figure 83: Avionics Bay Ventilated Temperature Sensor Location................................................ 21-185 Figure 84: Backup Exhaust and Air Extraction System Schematic ......................................................... 21-187 Figure 85: Avionics Exhaust Valve .................................... 21-189 Figure 86: Backup Fan....................................................... 21-191 Figure 87: Avionics Cooling and Extraction System Controls and Indications ................................... 21-193 Figure 88: Wing-to-Body Fairing Ventilation System ......... 21-199
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-vi
CS130_M5_21_EnviroContLOF.fm
21 - Environmental Control
Figure 89: High-Pressure Pneumatic Duct Area Cooling ..............................................................21-201 Figure 90: Cabin Pressure Control System ........................21-203 Figure 91: Typical Flight Profile ..........................................21-205 Figure 92: Pressurization Components Location................21-207 Figure 93: Outflow Valve ....................................................21-209 Figure 94: Outflow Valve, Water Activated Check Valve, and Muffler ........................................................21-211 Figure 95: Safety Valves and Negative Pressure-Relief Valves......................................21-213 Figure 96: Pressure Equalization Valves............................21-215 Figure 97: Integrated Air System Controller .......................21-217 Figure 98: Pressurization Panel .........................................21-219 Figure 99: FMS, Control Tuning Panel, and AVIONIC CTP Controls ..............................21-221 Figure 100:EICAS and AIR Synoptic Page Indications ......21-223 Figure 101:Cabin Pressure Control System Selection .......21-225 Figure 102:Cabin Pressure Control System Detailed Description ..........................................21-227 Figure 103:Cabin Altitude Limitations Annunciation ...........21-229 Figure 104:Emergency Depressurization ...........................21-231 Figure 105:Ditching.............................................................21-233
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_M5_21_EnviroContLOF.fm
ENVIRONMENTAL CONTROL - CHAPTER BREAKDOWN ENVIRONMENTAL CONTROL CHAPTER BREAKDOWN
Air Conditioning
1 Avionics Cooling and Extraction System
2 Cabin Pressure Control System
3
21 - Environmental Control 21-90 Integrated Air System Controller
21-90 INTEGRATED AIR SYSTEM CONTROLLER GENERAL DESCRIPTION Control and monitoring of the environmental control system (ECS) is provided by two integrated air system controllers (IASCs). The environmental control system includes: •
Air conditioning system
•
Avionics cooling and extraction system
•
Cabin pressure control system
The IASCs provide system information to the engine indication and crew alerting system (EICAS), the AIR synoptic page, and the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-2
CS130_21_L2_IASC_General.fm
21 - Environmental Control 21-90 Integrated Air System Controller
Pneumatic System
BAS
BALODS
EICAS
IASC 1
SYN OMS
WAIS
Environmental Control System IASC 2 ACS
CPCS
ACES ACS CPCS IASC
Bleed Air Cold Air Avionics Cooling and Extraction System Air Conditioning System Cabin Pressure Control System Integrated Air System Controller
ACES
CS1_CS3_2190_004
LEGEND
Figure 1: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-3
CS130_21_L2_IASC_General.fm
21 - Environmental Control 21-90 Integrated Air System Controller
COMPONENT LOCATION The IASCs are located in the mid equipment bay on the lower forward shelf.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-4
CS130_21_L2_IASC_ComponentLocation.fm
21 - Environmental Control 21-90 Integrated Air System Controller
IASC 2
IASC 1
INTEGRATED AIR SYSTEM CONTROLLER
CS1_CS3_2190_002
MID EQUIPMENT BAY FWD RACK
Figure 2: Integrated Air System Controller Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21-5
CS130_21_L2_IASC_ComponentLocation.fm
21 - Environmental Control 21-90 Integrated Air System Controller
DETAILED DESCRIPTION The IASC is a fully redundant controller. In case of failure of one channel, the other channel can control the ECS without loss of any function. The AIR SYSTEM FAULT advisory message is displayed for the failure of a channel. The IASC safety channel does an additional monitoring of all the most critical functions of the air systems. In the cooling and temperature control system, it monitors for overheat on temperature sensors and reports to the data concentrator unit module cabinet (DMC). In the avionics cooling and extraction system, it controls the extraction fans (EFAN) for the air extraction system, and reports to the DMC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_IASC_DetailedDescription.fm
21 - Environmental Control 21-90 Integrated Air System Controller
LEGEND IASC 1
AEV CSOV DMC
Channel A
Channel B
Safety Channel C
• L pack discharge disconnect and monitoring • Aft CSOV control and monitoring • Fwd galley heater temperature monitoring • EFAN no. 1 speed monitoring and control on OMS • Mid equipment bay temp monitoring • Control of both fwd/mid AEV and bufan • Fwd inlet fan control and monitoring
• L pack discharge disconnect and monitoring • R fan control and monitoring • Aft CSOV control and monitoring • Fwd galley heater temperature monitoring • Fwd equipment bay temperature monitoring • Fwd inlet fan control and monitoring
• EFAN no. 1 operation selection and failure detection
Data Exchange
EFAN IASC OMS
Avionics Exhaust Valve Cargo Shutoff Valve Data Concentrator Module Cabinet Exhaust Fan Integrated Air System Controller Onboard Maintenance System ARINC 429
DMC
IASC 2
Channel B
Safety Channel C
• R pack discharge disconnect and monitoring • Fwd CSOV control and monitoring • Aft galley heater temperature monitoring • EFAN no. 2 speed monitoring and monitoring report on OMS • Fwd equipment bay temperature monitoring • Control of both fwd/mid AEV and bufan • Fwd inlet fan control and monitoring • Mid ground valve control and monitoring
• R pack discharge disconnect and monitoring • R fan control and monitoring • Fwd CSOV control and monitoring • Aft galley heater temperature monitoring • Fwd equipment bay temperature monitoring • Mid inlet fan control and monitoring • Mid ground valve control and monitoring
• EFAN no. 2 operation selection and failure detection
CS1_CS3_2190_006
Channel A
Figure 3: Integrated Air System Controller Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_IASC_DetailedDescription.fm
21 - Environmental Control 21-90 Integrated Air System Controller
MONITORING AND TESTS The following page provides the INFO messages associated with the integrated air system controller:
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21-8
CS130_21_L2_IASC_MonitoringTests.fm
21 - Environmental Control 21-90 Integrated Air System Controller
INFO MESSAGES Table 1: INFO Messages MESSAGE
LOGIC
21 AIR SYSTEM FAULT - IASC 1A INOP
No data received from IASC1A confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 1B INOP
No data received from IASC1B confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 2A INOP
No data received from IASC2B confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 2B INOP
No data received from IASC2B confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 1C INOP
No data received from IASC1C confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - IASC 2C INOP
No data received from IASC2C confirmed 30s and channel power supply is available.
21 AIR SYSTEM FAULT - L IASC ARINC INPUT LOSS
No data from DMC1 or from DMC2 received by IASC1.
21 AIR SYSTEM FAULT - R IASC ARINC INPUT LOSS
No data from DMC1 or from DMC2 received by IASC2.
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CS130_21_L2_IASC_MonitoringTests.fm
21 - Environmental Control 21-90 Integrated Air System Controller
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM TEST The electrical environmental system components can be tested using the IBIT-ECS test on the onboard maintenance system (OMS). Follow the preconditions before starting the test.
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CS130_21_L2_IASC_PracticalAspects.fm
21 - Environmental Control 21-90 Integrated Air System Controller
MAINT MENU
RETURN TO FAULT MSGS
IASC1 chB-IBIT ECS
001 /004 IBIT ECS
EQUIPMENT TESTED: FCV, TASOV, TAPRV, TAV, DTS, VENTS, TCV, PACKS SENSORS, RARV
PRECONDITIONS
1 2 3 4 5
-
MAKE MAKE MAKE MAKE MAKE
SURE SURE SURE SURE SURE
WEIGHT ON WHEEL L BLEED SW IS SELECTED OFF R BLEED SW IS SELECTED OFF ALL ENGINE ARE OFF APU BLEED SW IS SELECTED OFF TEST INHIBIT/INTERLOCK CONDITIONS
WEIGHT ON WHEEL LEFT BLEED SWITCH
YYY
RIGHT BLEED SWITCH
YYY
ENGINE STATUS
YYY
APU BLEED SWITCH
YYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2190_003
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 4: Environmental Control System OMS Page Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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21-11
CS130_21_L2_IASC_PracticalAspects.fm
21 - Environmental Control 21-53 Flow Control System
21-53 FLOW CONTROL SYSTEM GENERAL DESCRIPTION The flow control system controls the airflow, airflow rate, and the temperature of bleed air provided into each air conditioning unit. The flow control system is installed on the system feed line, upstream of the air conditioning units. A trim air system is supplied with bleed air downstream of the flow control system. There is one system associated with each of the air conditioning units, installed on the aircraft. Each flow control system provides: •
Measurement and control of the air flow supplied to the air conditioning packs
•
Pack shutoff control
•
Pack overheat protection
An optional ozone converter may be installed in the bleed air ducting, upstream of the flow sensing venturi. The ozone converters convert ozone concentrations into oxygen, which prevents high ozone concentrations in the cabin. This improves the quality of air for high altitude or long duration flights. A second option adds a combined ozone and volatile organic compounds (VOC) converter.
The FCV is powered and controlled by the IASC. The left IASC controls the left FCV and the right IASC controls the right FCV. The IASC positions the FCV butterfly, using a torque motor, to limit the pressure and temperature of the air entering the pack. The FCV is spring-loaded closed. In the event of a power loss, the valve closes. A minimum 15 psi pressure is required to drive the valve in the open position. The maximum regulated pressure is limited to 43.5 +/- 3 psi. Each FCV is controlled by the associated PACK PBA on the AIR panel in the flight deck. The FCV is normally automatically controlled by the IASC but can be selected closed the L or R PACK PBA. During flight with reduced passenger loads, the system reduces the flow schedule in order to reduce bleed air consumption. Using the FMS data to capture passenger load, the IASC reduces the flow schedule based on the actual quantity of passengers onboard. The PACK FLOW PBA on the AIR panel, can be used to adjust the flow control to the high flow mode, which is used for smoke and odor clearance as well as increased pack performance.
The mass airflow into the air conditioning pack is measured by a flow sensor venturi (FSV) and associated sensors. The FSV is used to calibrate and measure the flow into the pack and trim air system. A pack inlet flow sensor (PIFS) provides differential pressure measurement across the flow sensor venturi to the IASC for mass flow computation and flow control valve (FCV) control. The pack inlet pressure sensor (PIPS) provides pack inlet static pressure to the IASC for mass flow computation and bleed air manifold pressure monitoring. The bleed air manifold pressure is displayed on the AIR synoptic page.
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21-12
CS130_21_L2_FCS_General.fm
21 - Environmental Control 21-53 Flow Control System
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
OFF
45
OPEN
XBLEED
PSI
45 PSI
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL
FMS
FAIL
SYNOPTIC PAGE – AIR
OFF R BLEED
DMC 1 DMC 2
FAIL
APU BLEED
IASC 1
OFF
FAIL
VOC
Volatile Organic Compounds
T/M
FSV
FCV
LEFT AIR CONDITIONING PACK
To Trim Air System
CS1_CS3_2150_013
DMC FCV FMS FSV IASC PIFS PIPS
ARINC 429 Data Concentrator Unit Module Cabinet Flow Control Valve Flight Management System Flow Sensor Venturi Integrated Air System Controller Pack Inlet Flow Sensor Pack Inlet Pressure Sensor
OZONE AND VOC CONVERTER (OPTIONAL)
L PIFS 2
LEGEND
Hot Air From Bleed Air System
L PIPS 2
AIR PANEL
L PIPS 1
L PIFS 1
OFF
Figure 5: Flow Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-13
CS130_21_L2_FCS_General.fm
21 - Environmental Control 21-53 Flow Control System
COMPONENT LOCATION The flow control system has the following components installed in the wing-to-body fairing (WTBF): •
Flow sensor venturi (FSV)
•
Flow control valve (FCV)
•
Ozone converter or combined ozone and volatile organic compounds (VOC) converter (Optional)
•
Pack inlet pressure sensor (PIPS) (Refer to figure 7)
•
Pack inlet flow sensor (PIFS) (Refer to figure 7)
FLOW SENSOR VENTURI The flow sensor venturi is located in the WTBF. FLOW CONTROL VALVE The flow control valve is located in the WTBF. OZONE CONVERTER/OZONE AND VOC CONVERTER (OPTIONAL) The ozone converter is installed upstream of the flow sensor venturi.
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21-14
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
Right Ozone and VOC Converter
Right Flow Sensor Venturi
Right Flow Control Valve
Left Ozone and VOC Converter
LEGEND VOC
Left Flow Control Valve
Volatile Organic Compounds
CS1_CS3_2150_030
Left Flow Sensor Venturi
Figure 6: Flow Sensor Venturi, Flow Control Valve, and Ozone Converter Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-15
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
PACK INLET PRESSURE SENSORS The pack inlet pressure sensors (PIPSs) are mounted on the pressure bulkhead in the wing-to-body fairing (WTBF). PACK INLET FLOW SENSORS The pack inlet flow sensors (PIFSs) are mounted on the pressure bulkhead in the WTBF.
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CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
Right PIF 1 Sensor
Right PIF 2 Sensor
Right PIP 2 Sensor
Right PIP 1 Sensor
Left PIP 2 Sensor Left PIP 1 Sensor
Left PIF 1 Sensor
CS1_CS3_2150_005 50 005
Left PIF 2 Sensor
NOTE PIP/PIF SENSORS
Left side shown.
Figure 7: PIP Sensor and PIF Sensor Locations Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-17
CS130_21_L2_FCS_ComponentLocation.fm
21 - Environmental Control 21-53 Flow Control System
CONTROLS AND INDICATIONS CONTROLS The L and R PACK PBA are used to select the flow control valves closed. An OFF legend on the PACK PBA indicates the valve is closed. The amber FAIL light indicates a pack fault. The HI FLOW PBA is used to increase the pack flow manually.
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21-18
CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
AIR COCKPIT
C
CABIN
H
C
H
FWD
MAN TEMP
C
CARGO
VENT
OFF
LO HEAT
OVERHEAD PANEL
AFT
OFF
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
VENT
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
H
HI HEAT
ON
FAIL
AFT
FAIL OFF R BLEED FAIL
APU BLEED
OFF
OFF
AIR PANEL
CS1_CS3_2160_006
FAIL
Figure 8: Flow Control Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-19
CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
INDICATIONS A HI indication appears in the pack symbol on the AIR synoptic page indicating the packs are in HI FLOW mode. The bleed air manifold pressure is supplied by the PIPS.
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CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
STATUS
AIR IR
DOOR OR
ELECC
FLT CTR CTRL TRLL
FUEL FUE EL
HYD YD
AVIONIC IONIC
INFO
CB
FWD
CKPT 23 °C 22 °C
CARGO
LO
15°CC
10 °C °
AFT
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2 15 °CC
VALVE POSITION NORMAL OPERATION Symbol
15 °CC
FLOW LINES Symbol
Condition
Condition Normal flow Closed
L PACK
HI
No flow
RAM AIR
TRIM AIR
Open with flow
R PACK
HI
Abnormal flow (overheat, out of range pressure, leak)
Open with no flow
VALVE POSITION FAIL OPERATION Symbol
45 PSI
XBLEED
45 PSI
Condition Closed Open with flow Open with no flow
CS1_CS3_2120_007
Invalid
APU
SYNOPTIC PAGE – AIR
Figure 9: Flow Control Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-21
CS130_21_L2_FCS_ControlsIndications.fm
21 - Environmental Control 21-53 Flow Control System
DETAILED DESCRIPTION The main operating modes for the system are the following:
FLOW CONTROL NORMAL OPERATION The flow is automatically controlled by the integrated air system controller (IASC) depending on the system configuration according to preset flow schedules. The FCV opens when the IASC provides a signal to the torque motor on the FCV and there is a minimum of 15 psi available from the bleed air system. The IASCs use the actual flow measured at the flow sensing venturi and the bleed temperature sensor (BTS) to modulate the FCV to obtain the required flow schedule. During flight with reduced passenger loads, the system automatically reduces the flow schedule in order to reduce bleed air consumption. Using the flight management system (FMS) passenger load data, the IASC reduces the flow schedule based on the actual quantity of passengers onboard. In the auto mode, the bleed air system uses auxiliary power unit (APU) bleed on the ground and during takeoff. The system switches to engine bleed air based on aircraft configuration. For additional information, refer to ATA 36 - Pneumatic.
•
Two bleed sources dual-pack operation
•
Two bleed sources single-pack operation
•
One bleed source dual-pack operation
•
Trim air only during emergency ram air operation
Pull-Up and Pull-Down Sequence When the APU is supplying bleed air on the ground and the actual zone temperatures vary by more than 10°C from the required temperature, the system automatically goes into pull-up or pull-down mode. Pull-up or pull-down modes use an increased flow schedule. During pull-up mode, additional heating is supplied by the trim air system to improve the pull-up efficiency.
On the ground, in dual pack or single pack operation, the APU flow demand is set to 100%, and the flow control logic uses a flow schedule based on: •
Aircraft version (CS100 or CS300)
•
Altitude (ranging from -2,000 ft to 14,500 ft)
•
Outside air temperature (-50°C to 50°C) (-58°F to 122°F)
In flight, the APU flow demand is set to 100%, and the system regulates the FCV according to the engine flow schedule.
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21-22
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
From L Bleed Air System
L PIPS 1
L PIFS 1
Recirc Air
L BTS
L FCV T/M
OZC
L PACK
IASC 1
APU
M To Trim Air System
DMC
IASC 2
R PIPS 1
R PIFS 1
APU ECU
MIX MANIFOLD
LEGEND
T/M
R PACK
OZC R PIPS 2
From R Bleed Air System
FSV R PIFS 2
R BTS
R FCV
Recirc Air
APU BTS DMC ECU FCV FSV HPGC IASC OZC PIFS PIPS
Hot Air Cold Air Auxiliary Power Unit Bleed Temperature Sensor Data Concentrator Unit Module Cabinet Electronic Control Unit Flow Control Valve Flow Sensor Venturi High-Pressure Ground Connection Integrated Air System Controller Ozone Converter Pack Inlet Flow Sensor Pack Inlet Pressure Sensor ARINC 429
CS1_CS3_2150_010
L PIFS 2
HPGC
L PIPS 2
FSV
Figure 10: Flow Control Normal Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-23
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
FLOW CONTROL To perform the flow control, the IASCs use the actual flow measured by: •
Pack inlet pressure sensor (PIPS)
•
Pack inlet flow sensor (PIFS)
•
Bleed temperature sensor (BTS)
The flow reference is function of: •
Number of passengers
•
Number of bleed air users
•
Aircraft altitude
•
Compressor discharge temperature sensor value
•
Bleed source selection
The flow requirements take into account the total passenger loads as well as the three cockpit members and five flight attendants.
NOTE 1. In dual-pack operation, with FWD cargo heat in either the LO or HI setting, and the recirculation air on, the flow schedule on the right side is increased to account for the added FWD cargo trim air. When the recirculation air system is off, but the FWD cargo heating is on, then both left and right FCVs are supplied with the same flow. 2. During single pack operation, the aircraft is altitude limited. The loss of a PIPS, PIFS, or an IASC control channel results in a PACK FAULT advisory message. Normal operation continues with the remaining sensors or the working channel. The loss of both PIPs, PIFs, or IASC control channels for either FCV is indicated by a L or R PACK FAIL caution message and an L(R) PACK PBA FAIL light. The pack must be shutdown by selecting the PACK PBA to OFF. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-24
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
IASC 1 Aircraft Altitude Number of Bleed Users Number of Passengers
FLOW CONTROL VALVE Flow Scheduled Value
T/M
CDTS Value Bleed Source Selection + -
L PIPS 1 L PIPS 2 Flow Measurement L PIFS 1 L PIFS 2
BTS
LEGEND Bleed Air Bleed Temperature Sensor Compressor Discharge Temperature Sensor Integrated Air Systems Controller Pack Inlet Flow Sensor Pack Inlet Pressure Sensor
CS1_CS3_2150_011
BTS CDTS IASC PIFS PIPS
Figure 11: Flow Control Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-25
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
FLOW CONTROL MODES Flow Priority When the flow measured is less than 80% of the minimum required flow for 3 minutes in single pack operation, or 90% in dual-pack operation, the flow priority mode is enabled. The flow priority circuit monitors for a low flow supply condition based on operating mode and passenger load as input in the FMS. The corresponding pack temperature control valve (TCV) is commanded open by the IASC. This allows increased air flow either until the flow target is achieved, or until the flight deck and cabin zone temperatures reach 35°C (95°F). The ram air regulating valve (RARV) opens to maintain the pack temperature. In this mode of operation, the flight deck and cabin zone temperature control are not available as the pack outlet temperature may be higher than the selected value. If the IASC detects a low flow condition for at least 3 minutes, an engine indication and crew alerting system (EICAS) L or R PACK FAIL advisory message is displayed.
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21-26
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control
L PIPS 1
L PIFS 1
21-53 Flow Control System
L FCV
CDTS
OZC
Reheater T/M
TURB
COMP
PTS
Condenser PDTS
L PIPS 2
L PIFS 2
FSV
M
M
PDPS
Air Cycle Machine TCV AIR CONDITIONING PACK
RARV
IASC 1
FWD CABIN VENTS
IASC 2
LEGEND Data Concentrator Unit Module Cabinet Flow Control Valve Flight Management System Integrated Air System Controller Ozone Converter Pack Inlet Flow Sensor Pack Inlet Pressure Sensor Pack Temperature Sensor Ram Air Regulating Valve Temperature Control Valve Ventilated Temperature Sensor ARINC 429
DMC 1 and DMC 2
AFT CABIN VENTS
FLIGHT DECK VENTS
FMS
CS1_CS3_2150_018
DMC FCV FMS IASC OZC PIFS PIPS PTS RARV TCV VENTS
Figure 12: Flow Control Priority Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-27
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
Hi Flow
FLOW CONTROL SHUTOFF
Whenever the PACK FLOW PBA is selected on the AIR panel, the flow control system switches to the high flow mode. The high flow mode corresponds to the maximum allowable flow for the system and it overrides all other flow control schedules. This mode is used to help evacuate cabin odors, smoke, or to increase the system air conditioning performance. The high flow mode can exchange the cabin air in less than 5 minutes in either dual and single PACK operating modes.
The flow control valve is closed when: •
The associated PACK PBA is selected OFF
•
Engine start in flight
•
Engine start on ground plus 30 seconds
•
Ditching PBA pressed
The HI flow mode is also automatically enabled when the auxiliary power unit (APU) is operating on the ground or when the required mix manifold temperatures decrease below 12°C (54°F) or increase above 60°C (140°F).
•
Bleed leak in pack zone as reported by the bleed air leak detection and overheat system
In flight, if the EMER DEPRESS PBA is selected, or if the cabin altitude increases above 15,000 ft, the flow control automatically switches to high flow mode.
When the FCV is commanded open with maximum torque motor current and the PIPS pressure is at least 15 psi and a low flow has been detected, a L or R PACK FAIL caution message is displayed.
FLOW CONTROL VALVE FAILURE
The FCV is detected failed open when it is commanded closed and the corresponding flow computed by the IASC is greater than 30 lb/min for more than 30 seconds. The L or R PACK FAIL caution message is displayed. This message is inhibited during emergency ram air mode with trim air ON since the flow through the FSV includes both the FCV and trim air system flow, which can exceed 30 lb/min in this mode.
NOTE During single engine bleed with: Wing anti-ice on (WAIS) Approach Steep approach (CS100 only) The flow control is limited to the reference flow, and HI flow is not available.
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21-28
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control
L FCV
CDTS
OZC
Reheater
EEC FCV FSV IASC PDPS PDTS PIFS PIPS PTS RARV TCV
TURB
COMP
PDTS PDPS
M
TCV
AIR CONDITIONING PACK Loop A
Loop B
DMC
M Bleed Air Warm Air Cold Air RARV Water Air Cycle Machine Compressor Discharge Temperature Sensor Data Concentrator Unit Module Cabinet Electronic Engine Control Flow Control Valve Flow Sensor Venturi Integrated Air System Controller Pack Discharge Pressure Sensor Pack Discharge Temperature Sensor Pack Inlet Flow Sensor Pack Inlet Pressure Sensor Pack Temperature Sensor Ram Air Regulating Valve Temperature Control Valve ARINC 429 Bleed Air Detection Loop A Bleed Air Detection Loop B
Condenser
Air Cycle Machine
LEGEND
ACM CDTS
PTS
T/M
L PIPS 2
L PIFS 2
FSV
IASC 1
IASC 2
PRESSURIZATION MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK
AUTO PRESS
L BLEED
FAIL
FAIL
ON
MAN CLSD
OFF
DITCHING
L EEC and R EEC
PRESSURIZATION PANEL
OFF
R PACK XBLEED
FAIL
MAN
OPEN
DMC AUTO
MAN OPEN
FAIL OFF R BLEED
APU BLEED FAIL OFF
AIR PANEL
FAIL OFF
CS1_CS3_2150_012
L PIPS 1
L PIFS 1
21-53 Flow Control System
Figure 13: Flow Control Shutoff Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-29
CS130_21_L3_FCS_DetailedDescription.fm
21 - Environmental Control 21-53 Flow Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the flow control system:
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21-30
CS130_21_L2_FCS_MonitoringTests.fm
21 - Environmental Control 21-53 Flow Control System
CAS MESSAGES Table 5: INFO Messages
Table 2: CAUTION Message MESSAGE L/R PACK FAIL
LOGIC
MESSAGE
Left/right pack inoperative (loss of control or monitoring [CDTS or PDTS failure] or pack low flow detected).
Table 3: ADVISORY Message MESSAGE PACK FAULT
LOGIC Loss of redundant or non-critical function for the air conditioning system.
Table 4: STATUS Message MESSAGE PACK FLOW HI
LOGIC CABIN AIR FLOW selected HI and confirmed HI by the controllers.
LOGIC
21 L PACK FAIL L FLOW CTRL VLV INOP
Left FCV failed open.
21 R PACK FAIL R FLOW CTRL VLV INOP
Right FCV failed open.
21 AIR SYSTEM FAULT - L PACK PRESS SNSR REDUND LOSS
Left pack pressure sensor redundant loss (PIFS or PIPS).
21 AIR SYSTEM FAULT - R PACK PRESS SNSR REDUND LOSS
Right pack pressure sensor redundant loss (PIFS or PIPS).
36 L BLEED FAIL - Left PIPS 1 and PIPS 2 loss. L PACK INLET PRESS SNSR INOP 36 R BLEED FAIL - R PACK INLET PRESS SNSR INOP
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TECHNICAL TRAINING MANUAL
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Right PIPS 1 and PIPS 2 loss.
21-31
CS130_21_L2_FCS_MonitoringTests.fm
21 - Environmental Control 21-53 Flow Control System
PRACTICAL ASPECTS The flow control valve can be dispatched in the closed position by capping and stowing the electrical connector and then removing the deactivation screw and installing it in the manual position indicator. If the valve is open, the manual position indicator can be used to close the valve using a wrench.
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21-32
CS130_21_L2_FCS_PracticalAspects.fm
21 - Environmental Control 21-53 Flow Control System
Visual Indicator and Lockout Mechanism
Deactivation Screw Position for Valve Lockout
CS1_CS3_2150_006 006
Deactivation Screw
Figure 14: Flow Control Valve Lockout Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_FCS_PracticalAspects.fm
21 - Environmental Control 21-51 Cooling
21-51 COOLING
ACM compresses and heats the air. The air passes through the main heat exchanger for further cooling before entering the reheater. The reheater heats the air entering the turbine inlet, where the remaining water is vaporized in order to prevent damage to the ACM.
GENERAL DESCRIPTION Cooling air is provided from two sources: •
Air conditioning units
•
Low-pressure ground cart
The air conditioning units, or packs are available on the ground or in flight. Each pack consists of: •
Air cycle machine
•
Dual-heat exchanger
•
High-pressure water extraction system
•
Temperature control valve
Air exiting the water extractor enters the turbine inlet where it expands and cools. This air is fed to the mix manifold for distribution to the aircraft.
The air cycle machine (ACM) consists of a turbine, compressor, and a fan connected by a common shaft. The ACM shaft is supported by air bearings. The dual-heat exchanger consists of a primary heat exchanger used to cool air from the bleed air system, and a main heat exchanger used to cool air from the air cycle machine (ACM) compressor discharge. The high-pressure water extraction system consists of a reheater, condenser, and a water separator. The extracted water is sprayed across the heat exchanger inlet to improve the cooling performance. The temperature control valve (TCV) bypasses bleed air around the ACM and exhausts it at the turbine outlet. This controls the pack outlet temperature and prevents pack icing. Each pack is controlled by an integrated air system controller (IASC). IASC 1 controls the left pack, and IASC 2 controls the right pack. Bleed air enters the pack through the flow control system. The bleed air is cooled by the primary heat exchanger before entering the ACM. The Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The air in the reheater starts cooling and any water vapor starts condensing. This air enters the condenser where it is cooled by the cold turbine outlet air. The water and air pass through the water extractor where the water is slung against the walls and is discharged through a line that runs to a spray tube upstream of the dual-heat exchanger. The water sprays across the face of the heat exchanger, increasing its cooling performance.
The water extraction line is monitored by the pack temperature sensor (PTS). If the air is too cold, there is a possibility of ice forming and damaging the ACM. To prevent this, the temperature control valve (TCV) is opened to allow some hot air to bypass the ACM, raising the temperature of the air. The dual-heat exchanger is cooled by ram air in flight. On the ground, the ACM fan draws air across the heat exchanger when the ACM is operating. The amount of cooling air drawn across the heat exchanger depends on the position of the ram air regulating valve (RARV). The pack temperature is controlled by modulating the RARV and the TCV. Pack overheat protection is provided by the compressor discharge temperature sensor and the pack discharge temperature sensor. If either sensor detects an overheat condition, a caution message is displayed and the pack is turned off using the PACK PBA which closes the flow control valve. On the ground, the low-pressure ground connection allows the use of an external ground cart to cool the aircraft. The low-pressure ground connection allows conditioned air to enter the left pack discharge duct downstream of the pack.
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CS130_21_L2_Cooling_General.fm
21 - Environmental Control 21-51 Cooling
From Bleed Air System
Ram Air Inlet
To Trim Air System Flow Control Valve
T/M
CDTS PTS
MIXTS
TURB
COMP
Reheater
PDTS
Condenser
To Mix Manifold
Air Cycle Machine M
M Temperature Control Valve
Ram Air Outlet
Water Extractor PDPS
Ram Air Regulating Valve
LEGEND
CS1_CS3_2160_003
CDTS MIXTS PDPS PDTS PTS
Bleed Air Warm Air Cold Air Water Compressor Discharge Temperature Sensor Mix Manifold Temperature Sensor Pack Discharge Pressure Sensor Pack Discharge Temperature Sensor Pack Temperature Sensor
Figure 15: Air Conditioning System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-35
CS130_21_L2_Cooling_General.fm
21 - Environmental Control 21-51 Cooling
COMPONENT LOCATION The following cooling system components are located in the wing-to-body fairing (WTBF): •
Air cycle machine
•
Dual-heat exchanger
•
Reheater condenser
•
Temperature control valve
•
Compressor discharge temperature sensor (CDTS)
•
Pack temperature sensor (PTS)
•
Plenum
•
Pack discharge temperature sensor (PDTS)
•
Ram air regulating valve (RARV)
•
RAM air duct
•
Water sprayer (Refer to figure 17)
•
Water extractor (Refer to figure 17)
The following cooling system component is located in the environmental control system bay: •
Pack discharge pressure sensor (Refer to figure 18)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
Pack Discharge Temperature Sensor
Water Extractor
Pack Temperature Sensor
Condenser/Reheater
Compressor Discharge Temperature Sensor Air Cycle Machine
Temperature Control Valve Plenum Dual Heat Exchanger
CS1_CS3_2150_019 S1 CS3 2150 019
Ram Air Duct
Figure 16: Air Conditioning Components Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-37
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
WATER EXTRACTOR The water extractor is mounted in front of the condenser/reheater. WATER SPRAYER The water sprayer is located in the ram air duct forward of the dual-heat exchanger. The sprayer discharge is oriented forward to optimize distribution across the face of the heat exchanger.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-38
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
Water Sprayer
Heat Exchanger Inlet
Water Extractor Water Extractor Drain
CS1_CS3_2150_007
Condenser/Reheater
Figure 17: Water Extractor and Water Sprayer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-39
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
PACK DISCHARGE PRESSURE SENSOR The pack discharge pressure sensor is located in the environmental control system bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-40
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
Pack Discharge Pressure Sensor Line
CS1_CS3_2150_020
Pack Discharge Pressure Sensor
Figure 18: Pack Discharge Pressure Sensor Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-41
CS130_21_L2_Cooling_ComponentLocation.fm
21 - Environmental Control 21-51 Cooling
COMPONENT INFORMATION TEMPERATURE CONTROL VALVE The temperature control valve (TCV) controls the pack outlet temperature, by extracting hot air downstream from the primary heat exchanger and reinjecting it at the turbine. The TCV is driven by a stepper motor. Microswitches provide position feedback to the IASC. In case of a power loss, the TCV remains in the last position. The TCV has a manual override lever to open or close the valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-42
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
Position Indicator
CS1_CS3_2160_013
Manual Override Lever
Figure 19: Temperature Control Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-43
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
RAM AIR REGULATING VALVE The ram air regulating valve (RARV) is used to regulate the ram air flow used on cold side of the PACK dual-heat exchanger. The RARV is driven by a stepper motor. Microswitches provide open and close position feedback. There is always flow through the valve, even when it is closed due to the fact that the butterfly valve is smaller than the valve housing. In case of power loss on the actuator, the RARV remains in its last position. The RARV has a manual override to open the valve for dispatch.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-44
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
Manual Lever
CS1_CS3_2150_027
Ram Air Regulating Valve
Figure 20: Ram Air Regulating Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-45
CS130_21_L2_Cooling_ComponentInformation.fm
21 - Environmental Control 21-51 Cooling
CONTROLS AND INDICATIONS CONTROLS The PACK PBA turns off the pack. An OFF legend on the PACK PBA indicates the pack is off. The PACK PBA FAIL legend indicates a fault has occurred that requires the pack to be shutdown.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-46
CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
AIR COCKPIT
C
FWD
H
CABIN
C
AFT
H
FWD
MAN TEMP
CARGO
VENT
OFF
C
LO HEAT
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK
OPEN
R PACK XBLEED
FAIL
MAN CLSD
OFF
AUTO
MAN OPEN
L BLEED
OFF
OFF R BLEED FAIL
APU BLEED
OFF
FAIL OFF
AIR PANEL
CS1_CS3_2150_025
FAIL
FAIL
Figure 21: Air Conditioning Pack Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-47
CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
INDICATIONS The pack status is shown on the AIR synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°C
10 °C
15 °C
15 °C RAM AIR
TRIM AIR
PACKS Symbol
R PACK
45 PSI
XBLEED
45
Condition
L PACK
Fail or leaked
L PACK
Active
L PACK
Invalid
L PACK
Normal OFF
PSI
CS1_CS3_2150_026
L PACK
APU
SYNOPTIC PAGE – AIR
Figure 22: Air Conditioning Pack Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_Cooling_ControlsIndications.fm
21 - Environmental Control 21-51 Cooling
DETAILED DESCRIPTION PACK TEMPERATURE REGULATION The pack temperature control is performed by two regulation loops within the IASC. The system controls the RARV and TCV simultaneously. The TCV is kept as close to the fully closed position as possible, minimizing the ram air flow, and the resulting induced drag penalty. On the ground, with the PACK selected ON, the RARV is commanded fully open above -19°C. Below this temperature the normal RARV/TCV control applies. In flight, when the PACK is OFF, the RARV is in its closed position. When the PACK is selected ON, and the flaps or the landing gear are not fully retracted, the RARV is commanded fully open. The TCV regulation loop compares mix manifold temperature sensor (MIXTS) values or PDTS if MIXTS is not available to a pack reference temperature. The RARV regulation loop compares compressor discharge temperature sensor (CDTS) value to a CDTS reference function of pack reference temperature. When the TCV or RARV fails, pack temperature control is maintained with the remaining TCV or RARV. An L or R PACK FAIL advisory message is displayed on EICAS for a TCV or RARV failure. PACK ICING PROTECTION To prevent ice buildup in the pack discharge outlet, the IASC opens the temperature control valve (TCV) to melt the ice. Icing protection is only available below 30,000 ft when the MIXTS is below 15°C (59°F). The ice formation is detected by the pack discharge pressure sensor (PDPS). The PDPS senses the difference between the pack discharge outlet and the cabin pressure .
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The TCV provides de-icing when the PDPS senses the differential pressure in the distribution duct is greater than 1.59 psi for more than 5 seconds. This mode is deactivated when the PDPS is less than 0.72 psi or after 30 seconds. In the event of a total loss of the PDPS, the pack discharge temperature is limited to 5°C (41°F) to prevent the formation of ice. PACK OVERHEAT PROTECTION The IASCs monitor the packs for an overheat condition using a compressor discharge temperature sensor (CDTS). The CDTS supplies an output to each IASC. The IACS safety channels monitor the CDTS and if the temperature exceeds 232°C (450°F) for 30 seconds or 260°C (500°F) for 5 seconds, an L or R PACK OVHT caution message is displayed. The pack must be shutdown manually using the PACK PBA. Selecting the PACK PBA FAIL light to OFF closes the flow control valve (FCV). To prevent a pack overheat condition, the system reduces the FCV flow as soon as the CDTS exceeds 220°C (428°F) and maintains the reduced flow until the temperature drops below this value. If the CDTS fails completely, the pack has no overheat protection and the same L or R PACK FAIL caution message is displayed. PACK DISCHARGE OVERHEAT PROTECTION In the normal mode, the pack discharge temperature is limited to 70°C (158°F). The pack discharge is monitored by the pack discharge temperature sensor (PDTS). An overheat occurs when the pack discharge temperature exceeds 85°C (185°F) for 30 seconds. When IASC safety channel detects a pack discharge overheat, a L or R PACK OVHT caution message is displayed. The FCV must be closed by selecting the PACK PBA FAIL light to OFF. If the PDTS fails completely, pack discharge overheat protection is not available, a L or R PACK FAIL caution message is displayed.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L3_Cooling_DetailedDescription.fm
21 - Environmental Control 21-51 Cooling
LEGEND Ram Air Inlet
FCV CDTS
TURB
Ram Air Regulating Valve
PTS
Air Cycle Machine
M
M
To Mix Manifold
Condenser PDTS
COMP
Reheater
MIXTS
T/M
From Bleed Air System
CDTS DMC FCV MIXTS PDPS PDTS PTS TCV
Bleed Air Cold Air Warm Air Water Compressor Discharge Temperature Sensor Data Concentrator Unit Module Cabinet Flow Control Valve Mix Manifold Temperature Sensor Pack Discharge Pressure Sensor Pack Discharge Temperature Sensor Pack Temperature Sensor Temperature Control Valve ARINC 429
TCV
Channel B
AIR COCKPIT
C
FWD
H
C
OFF
CABIN
H
FWD
MAN TEMP
VENT
IASC 1 Channel B
AFT
C
CARGO LO HEAT
Channel A
H
AFT
OFF
VENT
Safety Channel
DMC 1 and DMC 2
Channel A
Safety Channel
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
OPEN
CS1_CS3_2150_015
IASC 2
PDPS
Pressurized Cabin Underfloor
Figure 23: Air Conditioning System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-51
CS130_21_L3_Cooling_DetailedDescription.fm
21 - Environmental Control 21-51 Cooling
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the air conditioning system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-52
CS130_21_L2_Cooling_MonitoringTests.fm
21 - Environmental Control 21-51 Cooling
CAS MESSAGES
Table 9: INFO Messages Table 6: CAUTION Messages
MESSAGE L/R PACK FAIL
MESSAGE
LOGIC Left pack inoperative (loss of control or monitoring). Loss of L PACK leak detection capability. Loss of both IASC1 channels A and B.
L/R PACK FAIL
Right pack inoperative (loss of control or monitoring). Loss of R PACK leak detection capability. Loss of both IASC2 channels A and B.
L PACK OVHT
Left compressor discharge overheat or left pack discharge overheat.
R PACK OVHT
Right compressor discharge overheat or left pack discharge overheat.
Table 7: ADVISORY Message MESSAGE PACK FAULT
LOGIC Loss of redundant or non-critical function for the air conditioning system.
Table 8: STATUS Message MESSAGE L/R PACK OFF
LOGIC Left/right pack selected OFF and confirmed OFF.
Table 9: INFO Messages MESSAGE 21 PACK FAULT MIX MANF TEMP SNSR TOTAL LOSS
LOGIC MMTS from both packs are lost.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
21 PACK FAULT MIX MANF TEMP SNSR REDUND LOSS
Up to 3 MMTS sensing elements redundant loss or any sensing element drift.
21 PACK FAULT L BYPASS VLV INOP
Left TCV failed in position or position unknown.
21 PACK FAULT R BYPASS VLV INOP
Right TCV failed in position or position unknown.
21 L PACK FAIL L FLOW CTRL VLV INOP
Left FCV failed open.
21 R PACK FAIL R FLOW CTRL VLV INOP
Right FCV failed open.
21 L PACK FAIL L PACK INOP
L Pack failed for any reason except FCV failed open case.
21 R PACK FAIL R PACK INOP
R Pack failed for any reason except FCV failed open case.
21 PACK FAULT L PACK TEMP SNSR REDUND LOSS
Left CDTS or PDTS drift or single-channel failure.
21 PACK FAULT R PACK TEMP SNSR REDUND LOSS
Right CDTS or PDTS drift or single-channel failure.
21 PACK FAULT L PACK DISCH PRESS SNSR INOP
Left PDPS loss.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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21 - Environmental Control 21-51 Cooling
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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21 - Environmental Control 21-51 Cooling
Table 9: INFO Messages MESSAGE 21 PACK FAULT R PACK DISCH PRESS SNSR INOP
Table 9: INFO Messages
LOGIC
MESSAGE
Right PDPS loss.
21 L PACK OVHT - L PACK INOP
LOGIC L pack overheat triggered for any reason except RARV failed in position case.
21 R PACK OVHT R pack overheat triggered for any reason except RARV - R PACK INOP failed in position case.
36 L BLEED FAIL - Left PIPS 1 and PIPS 2 loss. L PACK INLET PRESS SNSR INOP
21 R PACK OVHT Left PTS drift or both channels failure. - L PACK TEMP SNSR INOP
36 R BLEED FAIL - R PACK INLET PRESS SNSR INOP
Right PIPS 1 and PIPS 2 loss.
21 R PACK OVHT Right PTS drift or both channels failure. - R PACK TEMP SNSR INOP L pack discharge duct disconnected.
21 AIR SYSTEM FAULT - L PACK PRESS SNSR REDUND LOSS
Left pack pressure sensor redundant loss (PIFS or PIPS).
21 L PACK FAIL L PACK DISCHARGE DUCT DISCONNECT
21 AIR SYSTEM FAULT - R PACK PRESS SNSR REDUND LOSS
Right pack pressure sensor redundant loss (PIFS or PIPS).
21 R PACK FAIL R PACK DISCHARGE DUCT DISCONNECT
R pack discharge duct disconnected.
21 PACK FAULT L RAM AIR REG VLV INOP
Left RARV failed in position or unknown position.
21 PACK FAULT R RAM AIR REG VLV INOP
Right RARV failed in position or unknown position.
21 L PACK OVHT Left RARV failed in position leading to pack overheat. - L RAM AIR REG VLV INOP 21 R PACK OVHT Right RARV failed in position leading to pack overheat. - R RAM AIR REG VLV INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_Cooling_MonitoringTests.fm
21 - Environmental Control 21-52 Ram Air System
21-52 RAM AIR SYSTEM GENERAL DESCRIPTION In case of dual pack loss, the emergency ram air valve provides a backup fresh air supply for unpressurized flight below 10,000 ft. The emergency ram air valve is powered by DC ESS BUS 3 and opens when the RAM AIR PBA is pressed. The emergency ram air valve connects the ram air inlet to the right pack discharge duct. If bleed air is available, the ram air can be mixed with hot air from the trim air system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-56
CS130_21_L2_RAS_General.fm
21 - Environmental Control 21-52 Ram Air System
DC ESS BUS 3
ICCP RAM AIR
OPEN
DMC 1 and DMC 2
RIGHT BULKHEAD CHECK VALVE Pack Discharge to Mix Manifold RIGHT PACK
M
EMERGENCY RAM AIR VALVE Ram Air Inlet
CS1_CS3_2150_023
LEGEND
Cold Air DMC Data Concentrator Unit Module Cabinet ICCP Integrated Cockpit Control Panel ARINC 429
Figure 24: Ram Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-57
CS130_21_L2_RAS_General.fm
21 - Environmental Control 21-52 Ram Air System
COMPONENT LOCATION The following components are located in the ram air system: •
Emergency ram air valve
EMERGENCY RAM AIR VALVE The emergency ram air valve is installed between the right pack discharge duct and the ram air inlet duct.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-58
CS130_21_L2_RAS_ComponentLocation.fm
21 - Environmental Control 21-52 Ram Air System
CS1_CS3_2120_017
Emergency Ram Air Valve
Figure 25: Emergency Ram Air Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-59
CS130_21_L2_RAS_ComponentLocation.fm
21 - Environmental Control 21-52 Ram Air System
CONTROLS AND INDICATIONS The RAM AIR PBA opens the emergency ram air valve. An OPEN legend on the RAM AIR PBA indicates the valve is open. The RAM AIR valve and flowbar on the AIR synoptic page illuminate green when the ram air system is selected ON.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-60
CS130_21_L2_RAS_ControlsIndications.fm
21 - Environmental Control 21-52 Ram Air System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
AIR COCKPIT
C
FWD
H
CABIN
C
AFT
H
FWD
MAN TEMP
CARGO
VENT
OFF
C
LO HEAT
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°C
10 °C
15 °C
H
15 °C RAM AIR
TRIM AIR
AFT
OFF
VENT
R PACK
L PACK
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
FAIL
PSI
XBLEED
45 PSI
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
45
OPEN
FAIL OFF R BLEED FAIL
APU BLEED
OFF
APU
OFF
AIR PANEL SYNOPTIC PAGE – AIR
CS1_CS3_2150_024
FAIL
Figure 26: Ram Air System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-61
CS130_21_L2_RAS_ControlsIndications.fm
21 - Environmental Control 21-52 Ram Air System
DETAILED DESCRIPTION The emergency ram air ventilation system supplies fresh airflow for unpressurized operation. The system supplies fresh air at altitudes up to 10,000 ft. The aircraft speed needs to be maintained above 0.5 Mach to ensure that enough dynamic air pressure is available at the ram air inlet. The emergency ram air valve connects the right hand pack ram air inlet diffuser to the right hand pack discharge duct. Air flows into the mix manifold for distribution throughout the aircraft. The valve has a manual override with a visual position indication. Microswitches provide position feedback to the data concentrator unit module cabinets (DMCs). The emergency ram air valve is powered from DC ESSENTIAL BUS 3 supplied through the hardwired RAM AIR guarded PBA on the AIR panel. An OPEN legend on the RAM AIR PBA, and a RAM AIR OPEN status message indicate the valve is open. When the RAM AIR PBA is selected, the outflow valve is commanded open to prevent excessive cabin pressure differential pressure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L3_RAS_DetailedDescription.fm
21 - Environmental Control 21-52 Ram Air System
ICCP Ram Air PBA
EMERGENCY RAM AIR VALVE Close
DC ESS BUS 3
3 Open Emer Ram Air Vlv Position
PIM
DMC 1 and DMC 2
Outflow Valve Open Command DMC ICCP PIM
CS1_CS3_2150_008
LEGEND Data Concentrator Unit Module Cabinet Integrated Cockpit Control Panel Panel Interface Module ARINC 429
Figure 27: Ram Air System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-63
CS130_21_L3_RAS_DetailedDescription.fm
21 - Environmental Control 21-52 Ram Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the ram air system:
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_RAS_MonitoringTests.fm
21 - Environmental Control 21-52 Ram Air System
CAS MESSAGES Table 10: CAUTION Message MESSAGE RAM AIR FAIL
LOGIC RAM AIR selected OPEN/CLOSED and ERAV not detected full open/full closed.
Table 11: STATUS Message MESSAGE RAM AIR OPEN
LOGIC RAM AIR selected and confirmed OPEN.
Table 12: INFO Messages MESSAGE
LOGIC
21 AIR SYSTEM ERAV in unknown position. FAULT - ERAV INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_RAS_MonitoringTests.fm
21 - Environmental Control 21-60 Trim Air System
21-60 TRIM AIR SYSTEM GENERAL DESCRIPTION The trim air system is used for independent zone temperature control. The aircraft zones serviced by the trim air are the flight deck, forward cabin, aft cabin, and forward cargo compartment. Bleed air is supplied from each pack inlet line upstream of the flow control valves. The trim air shutoff valve (TASOV) on the left side, and the trim air pressure-regulating valve (TAPRV) on the right side supply bleed air to the trim air system. The trim system takes hot air upstream the flow control valve and ducts it to the distribution system downstream of the mix manifold. The trim air is injected in the individual zone distribution air supply ducting, and mixed with cold air from packs to provide the required temperature for the zone. Each trim air line is fitted with a trim air check valve (TACKV), installed in line at the pressure bulkhead. In case of a trim air line disconnect it isolates the pressurized cabin from the unpressurized wing-to-body faring (WTBF) bay of the aircraft to prevent cabin depressurization.
In normal operation, the system is on and controlled automatically by the IASCs. The TAPRV modulates to provide 12.8 psi pressure, to the TAVs. The TASOV remains closed during normal operation. If the TAPRV fails closed, the TASOV is driven open by IASC 1 and used to supply the trim air system. To improve heating performance on the ground during the flow control pull up mode, the TASOV is opened by IASC 1. If the trim air system is selected off using the TRIM AIR PBA on the AIR panel or commanded closed when the flow control valves (FCVs) close, the TASOV, TAPRV, and TAVs close. The flight deck TAV remains slightly open to prevent a pressure build-up in the trim air distribution lines.
The TASOV and trim air valves (TAVs) are powered through their respective integrated air system controller (IASC). IASC 1 powers: •
TASOV
•
Flight deck TAV
•
Aft cabin TAV
IASC 2 powers: •
Forward cabin TAV
•
Forward cargo TAV
The TAPRV is powered by DC ESS BUS 1, and reports its position to IACSC 2. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-66
CS130_21_L2_TAS_General.fm
21 - Environmental Control 21-60 Trim Air System
LEGEND
CDC DMC
Bleed Air Cold Air Recirculation Air Conditioned Air Control and Distribution Cabinet Data Concentrator Unit Module Cabinet
FCV IASC
Flow Control Valve Integrated Air System Controller
TAPRV TASOV TAV
C
H
C
H
OFF
VENT
CARGO LO HEAT
OFF
HI HEAT
ON
IASC 1
Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve ARINC 429
C
FWD
MAN TEMP
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
IASC 2
DMC 1 and DMC 2
DC ESS BUS 1 CDC 2
LEFT PACK
From Bleed Air System
FLIGHT DECK
L FCV FWD CABIN EMERGENCY RAM AIR VALVE
RIGHT PACK
From Bleed Air System R FCV
MIX MANIFOLD
M
M
TASOV
FLIGHT DECK TAV
AFT CABIN
FWD CARGO COMPARTMENT M
M
TAPRV
AFT CABIN TAV
M FWD CARGO TAV
CS1_CS3_2160_016
FWD CABIN TAV SOL
Figure 28: Trim Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-67
CS130_21_L2_TAS_General.fm
21 - Environmental Control 21-60 Trim Air System
COMPONENT LOCATION The following components are located in the WTBF: •
Trim air pressure-regulating valve (TAPRV)
•
Trim air shutoff valve (TASOV)
•
Trim air check valve (TACKV)
The following components are located in the environmental control system bay: •
Trim air valves (TAVs) (Refer to figure 30)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-68
CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
Trim Air Check Valves
B
B TRIM AIR PRESSURE-REGULATING VALVE
A TRIM AIR SHUTOFF VALVE
CS1_CS3_2160_007
A
Figure 29: Trim Air Shutoff Valves, Trim Air Check Valves, and Trim Air Pressure-Regulating Valve Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
21-69
CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
TRIM AIR VALVES The trim air valves (TAVs) are located along the trim air distribution line in the environmental control system bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
Trim Air Distribution Line
Aft Cabin Trim Air Valve
Flight Deck Trim Air Valve
Fwd Cabin Trim Air Valve
TRIM AIR VALVE
CS1_CS3_2160_008
Cargo Trim Air Valve
Figure 30: Trim Air Valve Locations Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TAS_ComponentLocation.fm
21 - Environmental Control 21-60 Trim Air System
CONTROLS AND INDICATIONS The TRIM AIR PBA turns off the trim air system. An OFF legend on the TRIM AIR PBA indicates the trim air system is off. A TRIM AIR OFF message appears on the AIR synoptic page to indicate the system is off.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TAS_ControlsIndications.fm
21 - Environmental Control 21-60 Trim Air System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°C
10 °C
15 °C
15 °C
AIR COCKPIT
FWD
CABIN
AFT
R PACK
L PACK C
H
C
H
C
FWD
MAN TEMP OFF
VENT
CARGO LO HEAT
H
AFT
OFF
VENT
HI HEAT
OFF
45 RAM AIR TRIM AIR
RAM AIR
TRIM AIR
PACK FLOW
PSI
45
XBLEED
PSI
RECIRC AIR
OFF
APU
SYNOPTIC PAGE – AIR
CS1_CS3_2160_005
AIR PANEL
Figure 31: Trim Air System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TAS_ControlsIndications.fm
21 - Environmental Control 21-60 Trim Air System
DETAILED DESCRIPTION TRIM AIR CONTROL Each trim air valve (TAV) is commanded by the IASC according to the duct temperature sensor (DTS) measurement compared with a reference temperature, which is determined by the error between the selected temperature from the AIR panel in the flight deck and the temperature measured by the associated zone ventilated temperature sensor (VENTS). When the trim air system is selected OFF, the TAVs are also commanded closed by the IASCs. The flight deck TAV is commanded slightly open to avoid a pressure buildup in the trim air lines. The flight deck, forward cabin, aft cabin, and the FWD cargo TAV are automatically closed whenever the trim air system is closed, and either the TAPRV or the TASOV fails open.
•
Pack start up
•
Engine start + 30 seconds
•
Ditching
The cargo TAV is automatically closed when: •
The cargo heating system is selected OFF by the flight crew FWD CARGO switch set to OFF or VENT
•
The cargo SOVs are commanded closed due to forward cargo compartment smoke detection
•
TRIM AIR PBA OFF
•
Cargo DTS overheat
•
Total loss of cargo DTS
NOTE When the trim system is shutoff and all TAVs are closed, the flight deck TAV is commanded to open for 2.7 seconds, to prevent pressure build-up in the trim air distribution line. The TASOV and TAPRV are automatically closed whenever: •
DTS overheat is detected
•
Bleed leak is detected affecting the TAPRV and TASOV downstream ducting
•
Both flow control valves (FCVs) are commanded closed, and the emergency ram air valve (ERAV) is not selected open
•
TRIM AIR PBA is selected OFF
•
The TAPRV or the TASOV are commanded closed when its associated FCV is closed, and ERAV is not selected open.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-60 Trim Air System
LEGEND Bleed Air Cold Air Recirculation Air Conditioned Air Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Duct Temperature Sensor Integrated Air System Controller Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve Ventilated Temperature Sensor ARINC 429 Discrete
Recirculation Air
From Left Pack
MIXING MANIFOLD
From Right Pack
DTS
FLIGHT DECK VENTS
CDC DMC DTS IASC TAPRV TASOV TAV VENTS
Recirculation Air IASC 1
Channel B COCKPIT
From Bleed Air System From Bleed Air System
M
FLIGHT DECK TRIM AIR VALVE
M SOL
TAPRV
To Other Trim Air Valves
+ -
Flight Deck Duct Reference Channel A
NOTE
+
C
CABIN
H
C
CARGO
MAN TEMP
HEAT
-
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
AIR PANEL
Flight deck shown, FWD cabin, aft cabin, and FWD cargo similar.
AFT
OPEN
CS1_CS3_2150_017
TASOV
Figure 32: Trim Air Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-75
CS130_21_L3_TAS_DetailedDescription.fm
21 - Environmental Control 21-60 Trim Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the trim air system:
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TAS_MonitoringTests.fm
21 - Environmental Control 21-60 Trim Air System
CAS MESSAGES Table 16: INFO Messages
Table 13: CAUTION Messages MESSAGE
LOGIC
MESSAGE
LOGIC
TRIM AIR FAIL
Trim air failure or loss of trim air leak detection or loss of both IASC 1channels.
FWD CARGO HEAT FAIL
LO HEAT AND HI HEAT mode not available.
21 AIR SYSTEM TAPRV failed closed. FAULT - TRIM AIR PRV FAIL CLSD
FWD CARGO LO TEMP
Low temperature <4°C in fwd cargo when FWD CARGO selected to LO/HI HEAT
21 AIR SYSTEM TAPRV failed open. FAULT - TRIM AIR PRV FAIL OPEN 21 AIR SYSTEM TASOV failed closed in full closed position. FAULT - TRIM AIR SOV FAIL CLSD
Table 14: ADVISORY Message MESSAGE AIR SYSTEM FAULT
LOGIC Fault not requiring immediate attention. (Total loss of VENTS, IASC single-channel failure, loss of cockpit, FWD, AFT cabin DTS, TAV stuck, TAPRV/TASOV failed closed, single channel CAR DTS, or CATS failed).
Table 15: STATUS Messages MESSAGE TRIM AIR OFF
LOGIC TRIM AIR selected OFF and both TAPRV and TASOV confirmed closed.
FWD CARGO AIR FWD cargo heating selection selected and confirmed OFF OFF. AFT CARGO AIR OFF
AFT cargo heating selection selected and confirmed OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
21 AIR SYSTEM TASOV failed not closed or unknown position. FAULT - TRIM AIR SOV FAIL OPEN 21 AIR SYSTEM FAULT - FWD CARGO SOV INOP
Upstream or downstream fwd cargo SOV failed in position or in unknown position.
21 AIR SYSTEM FAULT - AFT CARGO SOV INOP
Upstream or downstream aft cargo SOV failed.
21 AIR SYSTEM FAULT - FWD CAR TAV FAIL CLSD
CAR TAV failed closed.
21 AIR SYSTEM Loss of CKPT or FWD CAB or AFT CAB TAV or any TAV FAULT - TAV INOP unknown position
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21 - Environmental Control 21-60 Temperature Control System
21-60 TEMPERATURE CONTROL SYSTEM GENERAL DESCRIPTION The temperature control system controls the flight deck, forward and aft cabin, and the forward cargo compartment. Individual automatic temperature controls for the flight deck and forward and aft cabins are available on the AIR panel. The forward and aft cabin temperature is also controlled from the cabin management system (CMS) crew terminal. The forward cargo heat is selectable from the AIR panel. The system temperature control for the flight deck, forward and aft cabins, and forward cargo compartment is automatically controlled via the integrated air system controllers (IASCs) in order to meet the flight deck temperature selections. The IASCs receive temperature inputs from: •
Ventilated temperature sensor for actual zone temperature
•
Duct temperature sensors for duct temperature
•
Mix manifold temperature sensors for pack temperature control
•
Cargo temperature sensor for forward cargo actual temperature
•
Cargo duct temperature sensor for cargo duct temperature
The forward and aft cabin temperature can be adjusted at the cabin management system crew terminal. The crew terminal provides the capability to adjust the temperature by ± 3°C (5.4°F) from the temperature selection made on the AIR panel FWD, or AFT CABIN temperature selectors. The forward cargo LO HEAT provides a temperature range of 15°C (59°F) to 20°C (68°F), and HI HEAT provides a temperature range of 20°C (68°F) to 25°C (77°F).
The packs are controlled to provide the lowest temperature output required by the flight deck, forward, or aft cabin zones. The IASCs control the trim air valve (TAV) to add hot air to each of the distribution ducts that require additional heating. In automatic mode, the required zone temperature for the flight deck, forward, and aft cabin can be set from 18°C (64°F) to 30°C (86°F) using the AIR panel temperature selectors. In manual mode, the flight crew adjust the zone temperatures by adjusting the duct temperature from full cold 5°C (41°F) to full hot 65°C (149°F) using the AIR panel temperature selectors.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-78
CS130_21_L2_TCS_General.fm
21 - Environmental Control 21-60 Temperature Control System
LEGEND Hot Air Cold Air Recirculation Air Conditioned Air Cargo Duct Temperature Sensor Cargo Temperature Sensor Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Duct Temperature Sensor Integrated Air System Controller Mix Manifold Temperature Sensor Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve Ventilated Temperature Sensor ARINC 429
C
FWD
H
C
H
FWD
MAN TEMP OFF
CABIN
VENT
AFT
C
CARGO LO HEAT
CABIN MANAGEMENT SYSTEM CREW TERMINAL
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
OPEN
DMC 1 and DMC 2
IASC1
IASC2
CDC 2
DTS
CADTS CATS CDC DMC DTS IASC MIXTS TAPRV TASOV TAV VENTS
AIR COCKPIT
From Left Pack
UNDERFLOOR CATS
DTS
MIX MANIFOLD
MIXTS 2
From Bleed Air System
TASOV M
TAPRV
SOL
VENTS
FWD CARGO COMPARTMENT
M AFT CABIN TAV FWD CARGO TAV M
FWD CABIN
AFT CABIN
M F/D TAV M FWD CABIN TAV
VENTS
CADTS
CS1_CS3_2160_021
From Right Pack
From Bleed Air System
FLIGHT DECK DTS
MIXTS 1
VENTS
Figure 33: Temperature Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-79
CS130_21_L2_TCS_General.fm
21 - Environmental Control 21-60 Temperature Control System
COMPONENT LOCATION The following are major components of the temperature control system: •
Mix manifold temperature sensors (MIXTS)
•
Duct temperature sensor (DTS) (Refer to figure 35)
•
Ventilated temperature sensors (VENTS) (Refer to figure 36)
•
Cargo duct temperature sensor (CADTS) (Refer to figure 37)
•
Cargo temperature sensor (CATS) (Refer to figure 37)
MIX MANIFOLD TEMPERATURE SENSORS The mix manifold temperature sensors are located in the pack discharge lines near the inlet to the mix manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
Mix Manifold Temperature Sensors
CS1_CS3_2120_029
TEMPERATURE SENSOR
Figure 34: Mix Manifold Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-81
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
DUCT TEMPERATURE SENSORS The duct temperature sensors (DTS) for the forward cabin, aft cabin, and flight deck are located in the environmental control system bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-82
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
Fwd Cabin Duct Temperature Sensor
Aft Cabin Duct Temperature Sensor TEMPERATURE SENSOR
CS1_CS3_2160_011
Flight Deck Duct Temperature Sensor
Figure 35: Duct Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-83
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
VENTILATED TEMPERATURE SENSORS Flight Deck Ventilated Temperature Sensor The flight deck ventilated temperature sensor is installed on the pilots bulkhead. Forward Cabin Ventilated Temperature Sensor The forward cabin ventilated temperature sensor is located on the right side of the forward cabin. The sensor is mounted in the passenger service unit located between frame 34 and frame 35. Aft Cabin Ventilated Temperature Sensor The aft cabin ventilated temperature sensor is located on the left side of the aft cabin. The sensor is mounted in the passenger service unit located between frame 62 and frame 64.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-60 Temperature Control System
C
A
B
A
VENTILATED TEMPERATURE SENSOR
Ventilated Temperature Sensor
B AFT CABIN
C FWD CABIN
CS1_CS3_2160_010
Ventilated Temperature Sensor
Figure 36: Flight Deck and Cabin Ventilated Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-85
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
CARGO DUCT TEMPERATURE SENSOR The cargo duct temperature sensor (CADTS) is located in the forward cargo ventilation inlet duct in the environmental control system (ECS) bay. CARGO TEMPERATURE SENSOR The cargo temperature sensor (CATS) is located in the forward cargo ventilation exhaust duct in the environmental control system bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
Fwd Cargo Compartment
TEMPERATURE SENSOR Cargo Temperature Sensor
CS1_CS3_2160_009
Cargo Duct Temperature Sensor
Figure 37: Cargo Duct Temperature Sensor and Cargo Temperature Sensor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-87
CS130_21_L2_TCS_ComponentLocation.fm
21 - Environmental Control 21-60 Temperature Control System
COMPONENT INFORMATION VENTILATED TEMPERATURE SENSOR The ventilated temperature sensor (VENTS) supply temperature information to the IASC. The vents consist of two temperature sensing elements installed on a circuit board. The circuit board is housed in a plastic box along with a small fan. The fan draws air across the temperature sensor. A filter is installed on the cover to ensure the temperature sensors stay clean.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ComponentInformation.fm
21 - Environmental Control 21-60 Temperature Control System
CS1_CS3_2160_022
Filter
Figure 38: Ventilated Temperature Sensor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-89
CS130_21_L2_TCS_ComponentInformation.fm
21 - Environmental Control 21-60 Temperature Control System
CONTROLS AND INDICATIONS AIR PANEL The COCKPIT, FWD and AFT CABIN temperature control selectors are used to set the desired temperature in automatic and manual modes. A MAN TEMP PBA is to be used to select manual control of the temperature control system. An ON legend indicates the system is in manual control. The FWD CARGO rotary switch is used to select LO HEAT or HI HEAT in the forward cargo compartment.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
AIR COCKPIT
C
FWD
H
CABIN
C
H
C
FWD
MAN TEMP
CARGO
VENT
OFF
AFT
LO HEAT
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR PACK FLOW
RECIRC AIR
L PACK
R PACK AIR PANEL
CS1_CS3_2160_020
TRIM AIR
Figure 39: Temperature System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-91
CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
AIR SYNOPTIC PAGE The AIR synoptic page indicates the actual temperature for the flight deck, forward, and aft cabins on top. The selected temperature from the AIR panel is shown below the actual temperature. If the selected temperature has been modified by the cabin management system (CMS) crew terminal temperature page selection, that information is displayed beside the selected temperature. The duct temperature is indicated at the outlet of the mix manifold representation on the synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
STATUSS
AIR IR
DOOR OR
ELECC
FLT CTR CTRLL
FUEL FUE EL
HYD YD
AVIONIC
INFO
CB
AIR SYNOPTIC PAGE INDICATIONS
2
3
CKPT 00 °C 24 °C
24°CC
1
CARGO
LO 15 °C °
FWD 26 °C4 24 °°C -+ 2
AFT 26 °C 24 °C -+ 2
24 °CC
Fwd Cargo Selection Indicator: • OFF – no heating, no ventilation • VENT – cargo ventilation only • LO HEAT – vent and low heating level (15-20°C) • HI HEAT – vent and warmer heating level (20-25°C)
1
CARGO LO
2
AFT 26°C
Actual Cabin Temperature: • Dashed line "ííí" represents a temperature sensor failure or operating in manual mode
3
24°C
Selected Temperature: • Automatic mode – desired zone temperature • Manual mode – desired duct temperature
4
+2
Cabin Management System Input (+/-3°C)
5
24°C
Duct Temperature
24 °C 5
RAM AIR
R PACK
L PACK
CS1_CS3_2160_018
SYNOPTIC PAGE – AIR
Figure 40: Temperature System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
CABIN MANAGEMENT SYSTEM TEMPERATURE CONTROLS The current temperature levels, and the temperature adjustment for each cabin zone are displayed on the temperature screen. The allowable range is displayed for each cabin zone. The current temperature in each cabin is displayed on a layout of passenger area (LOPA) screen. If the temperature adjustment is not available, the temperature controls are disabled and an error message is displayed on the screen.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control
CS1_CS3_2160_019
21-60 Temperature Control System
Figure 41: Cabin Management System Temperature Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L2_TCS_ControlsIndications.fm
21 - Environmental Control 21-60 Temperature Control System
DETAILED DESCRIPTION AUTOMATIC TEMPERATURE CONTROL The automatic temperature control system controls the air conditioning packs, ram air regulating valve (RARV), recirculation system, and the trim air system to provide the selected temperature for the flight deck, the forward, and aft cabins, and the forward cargo compartment. For zone temperature control, the integrated air system controllers (IASCs) use the following sensors: •
The ventilated temperature sensors (VENTS), located in the flight deck, forward cabin and aft cabin zones sense the actual zone temperature
•
The duct temperature sensors (DTS), located in the main zone distribution lines downstream of the TAV, sense the temperature of the air supplied to each zone
•
The mix temperature sensors (MIXTS) sense the temperature of the pack discharge and recirculation mixed air for the pack temperature control valve (TCV)
•
The pack discharge temperature sensors (PDTS), located on pack discharge ducts, sense each pack discharge temperature. The PDTS is mainly used as a monitoring means, such as pack discharge overheat protection or as a backup to the MIXTS, the PTS or the PDPS for redundant icing protection.
The IASC modulates the TCV and RARV position to achieve the reference mix manifold temperature. The mix manifold reference temperature is set equal to the lowest reference temperature demand computed for each of the three independent zones. The IASC controls each TAV in a closed loop against the duct temperature sensor measurement. The DTS reference temperature is computed from the difference between the ventilated temperature sensor measurement and the zone temperature selections made on the AIR panel.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Each TAV is commanded by its IASC according to the duct temperature sensor (DTS) measurement compared with a reference temperature determined by the error between the selected temperature from the AIR panel in the flight deck, and the temperature measured by the associated zone ventilated temperature sensor (VENTS). •
If the forward or aft cabin VENTS fails, the remaining VENTS is used for temperature control in both cabins
•
If the flight deck VENTS or both the forward and aft cabin VENTS fails, the affected zones are controlled using the supply duct temperature as a reference similar to the manual temperature control
•
If a temperature selector on the AIR panel fails, the system assumes a default value of 24°C (75°F) for that zone
•
If the forward or aft DTS fails, the associated TAV is closed and the remaining zone provides temperature control for the entire cabin
•
If the flight deck DTS fails, the TRIM AIR PBA must be selected OFF
If the flight deck, forward or aft cabin DTS measurement exceeds 85°C (185°F) for over 30 seconds, the trim air system is closed automatically, and a TRIM AIR FAIL caution message is posted and latched until the TRIM AIR PBA is selected OFF, and the overheat condition disappears. MANUAL TEMPERATURE CONTROL In manual mode, the flight crew can manually adjust the zone temperatures from the AIR panel by selecting the MAN TEMP PBA to ON. The duct temperature sensors (DTS) are used for the reference temperature. The temperature in the duct can be controlled from 5°C (41°F) to 65°C (149°F). When in manual temperature control, the selected temperature reading on the synoptic page becomes the DTS target, and the actual temperature displayed on the synoptic page becomes dashed.
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21 - Environmental Control 21-60 Temperature Control System
Ram Air Inlet
AIR COCKPIT
C
FWD
H
C
CABIN
H
FWD
MAN TEMP OFF
VENT
PTS
CDTS
H
AFT
OFF
VENT
HI HEAT
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
-
Hot Air Warm Air Cold Air Water Duct Temperature Sensor Integrated Air System Controller Minimum Mix Manifold Temperature Sensor Pack Discharge Temperature Sensor Pack Temperature Sensor Ram Air Regulating Valve Trim Air Valve Temperature Control Valve IASC Ventilated Temperature Sensor
+
VENTS
F/D TAV M
Temperature Control Valve
LEGEND
FLIGHT DECK
+
-
-
Flight Deck Duct Reference
Trim Reference
MIXTS Reference
+
Min Temp
+
FWD Cabin Duct Reference
-
Trim Air Available
+
Aft Cabin Duct Reference
-
FWD Cabin Vents
Aft Cabin Vents
CS1_CS3_2161_001
Ram Air Regulating Valve
MIX MANIFOLD
To PDPS
M
M
MIXTS
PDTS
Condenser
DTS
TURB
COMP
Reheater
Air Cycle Machine
DTS IASC MIN MIXTS PDTS PTS RARV TAV TCV VENTS
C
CARGO LO HEAT
ON
From Bleed Air System
AFT
Figure 42: Flight Deck, Forward Cabin, and Aft Cabin Temperature Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_TCS_DetailedDescription.fm
21 - Environmental Control 21-60 Temperature Control System
FWD CARGO TEMPERATURE CONTROL The TAV mixes hot air to the flow distributed from the recirculation system to the forward cargo in order to provide the required temperature. The temperature is controlled in a closed loop by measurement feedback from the cargo temperature sensor (CATS) to IASC 2. The cargo duct temperature sensor (CADTS) provides temperature measurement to IASC 2 for overheat monitoring and protection. A zone target temperature is set, based on LO or HI heat selection and the duct temperature reference, based on the error between this target and the measured temperature. If the cargo duct temperature exceeds 70°C (158°F) for 30 seconds, the trim air system is closed automatically. A TRIM AIR FAIL caution message is displayed until the TRIM AIR PBA is selected OFF, and the overheat condition disappears. If the CADTS or CATS fails, the FWD CARGO HEAT is selected OFF. When the FWD cargo selector is set to OFF or VENT, and the CARGO TAV full closed switch is not received, a TRIM AIR FAIL caution message is posted and latched until the TRIM AIR PBA is selected OFF.
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21 - Environmental Control 21-60 Temperature Control System
CATS DOWNSTREAM CARGO SOV FWD CARGO COMPARTMENT IASC2 UPSTREAM CARGO SOV
M From Trim Air System
+
FWD Cargo Duct Reference
CADTS
Recirculation Air
FWD + -
OFF
VENT
CA LO HEAT HI HEAT
AIR PANEL
FWD CARGO TAV
LEGEND
CS1_CS3_2161_002
CADTS CATS IASC SOV TAV
Hot Air Warm Air Cold Air Water Cargo Duct Temperature Sensor Cargo Temperature Sensor Integrated Air System Controller Shutoff valve Trim Air Valve
Figure 43: Cargo Temperature Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_21_L3_TCS_DetailedDescription.fm
21 - Environmental Control 21-60 Temperature Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the temperature control system:
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21 - Environmental Control 21-60 Temperature Control System
CAS MESSAGES Table 20: INFO Messages
Table 17: CAUTION Messages MESSAGE
LOGIC
MESSAGE
FWD HEAT CARGO FAIL
Total loss of CADTS or CATS.
FWD CARGO LO TEMP
In the event the FWD cargo is selected to LO or HI HEAT, and the CATS reads below 10°C for at least 2 minutes during cruise.
Table 18: ADVISORY Messages MESSAGE AIR SYSTEM FAULT
PACK FAULT
LOGIC Fault not requiring immediate attention. (Total loss of VENTS, IASC single channel failure, loss of cockpit, FWD, AFT cabin DTS, TAV stuck, TAPRV/TASOV failed closed, single-channel CADTS, or CATS failed). Loss of redundant or non-critical function for the air conditioning system.
LOGIC
21 AIR SYSTEM FAULT - ZONE TEMP SNSR INOP
Cockpit VENTS total loss (both channels) OR
21 AIR SYSTEM FAULT - DUCT TEMP SNSR INOP
Total loss or drift of FWD CABIN DTS or AFT CABIN DTS or redundant loss or drift of CKPT DTS.
21 PACK FAULT MIX MANF TEMP SNSR TOTAL LOSS
MMTS from both packs are lost.
21 PACK FAULT MIX MANF TEMP SNSR REDUND LOSS
Up to three MMTS sensing elements redundant loss.
Aft Cabin VENTS total loss (both channels) OR FWD Cabin VENTS total loss (both channels).
Table 19: STATUS Messages MESSAGE
LOGIC
FWD CARGO AIR FWD cargo heating selection selected and confirmed OFF OFF. AFT CARGO AIR OFF
AFT cargo heating selection selected and confirmed OFF.
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21 - Environmental Control 21-22 Recirculation
21-22 RECIRCULATION GENERAL DESCRIPTION Cabin air is recycled into the mixing manifold by mean of a single recirculation fan (RFAN), which draws air from the cabin underfloor across two air filters. The RFAN supplies the air into each pack discharge line for initial mixing with the pack discharge air, downstream of the bulkhead check valve. Integrated air system controller (IASC) 1 controls the RFAN. If IASC 1 fails, IASC 2 controls the recirculation fan. The recirculation fan has a motor controller that the IASC signals to adjust the fan speed based mainly on cabin altitude. The fan is powered by AC BUS 2, and the controller is powered by DC BUS 2. The recirculation fan operation is fully automatic and runs when power is available, no faults exist, and there is no smoke in the equipment bays. The recirculation fan can be selected OFF using the RECIRC AIR PBA on the AIR panel. The RFAN is normally selected OFF when the fire detection and extinguishing (FIDEX) control unit reports smoke in the equipment bays. The recirculation fan is commanded so that the recirculation flow ratio remains between 25% and 40% of the total air flow for normal operation, and between 35% and 50% of the total airflow for single pack operation. The recirculation flow also ensures the mixing manifolds temperature remains above freezing.
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21 - Environmental Control 21-22 Recirculation
From Left Hand Pack
From Right Hand Pack
Low Pressure Ground Connection
Emergency Ram Air Valve
Pressurized Cabin Underfloor
Pressurized Cabin Underfloor
Check Valves
Cargo Ventilation MIXING MANIFOLD
RECIRCULATION FILTER
RECIRCULATION FILTER
RECIRC AIR FIDEX CONTROL UNIT
DMC 1 and DMC 2
RECIRCULATION FAN
OFF
AC BUS 2 DC BUS 2
DMC IASC
Cold Air Recirculated Air Data Concentrator Unit Module Cabinet Integrated Air System Controller ARINC 429 CAN BUS
IASC 1 and IASC 2
CS1_CS3_2120_019
LEGEND
Figure 44: Recirculation System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
COMPONENT LOCATION The following components are located in the environmental control system bay, aft of the forward cargo compartment: •
Recirculation fan filters
•
Recirculation fan and motor controller
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21 - Environmental Control 21-22 Recirculation
CS1_CS3_2120_012
Recirculation Fan Filters
Recirculation Fan and Motor Controller
Figure 45: Recirculation System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-105
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21 - Environmental Control 21-22 Recirculation
COMPONENT INFORMATION RECIRCULATION FILTERS The disposable recirculation filter ensures cleaned air is recycled to the cabin. It is composed of an HEPA matrix, and an active charcoal core containing chemical absorbers used to control odors. The recirculation filter is supported by a structure at the bottom and at the top. A V-band clamp holds the filter at the bottom, and a plunger centers the filter at the top.
WARNING DO NOT TOUCH THE FILTER WITH BARE HANDS. USE GLOVES AND PERSONAL PROTECTIVE EQUIPMENT TO HANDLE THE FILTER. THE FILTER SHOULD BE CONSIDERED A BIOHAZARD. THE FILTER CAN CONTAIN GERMS AND VIRUSES THAT CAN CAUSE SICKNESS.
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21 - Environmental Control
CS1_CS3_2120_026
21-22 Recirculation
Figure 46: Recirculation Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
CONTROLS AND INDICATIONS The RECIRC AIR PBA turns off the recirculation fan. An OFF legend on the RECIRC AIR PBA indicates the system is shutdown. The RECIRC AIR PBA is used when there is a recirculation fan failure or smoke is detected in the avionics cooling air extraction system. The PBA is selected to match the system configuration. A RECIRC OFF message appears on the AIR synoptic page to indicate the system is shutdown.
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21-108
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21 - Environmental Control 21-22 Recirculation
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
FWD
AFT
CKPT 23 °C 22 °C
CARGO
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°CC
10 °°CC
15 °CC
15 °CC
AIR COCKPIT
FWD
CABIN
AFT
RAM AIR
TRIM AIR R PACK
L PACK C
H
C
H
C
FWD
MAN TEMP OFF
VENT
CARGO LO HEAT
RECIRC OFF
H
AFT
OFF
VENT
HI HEAT
45 RAM AIR TRIM AIR
PACK FLOW
PSI
XBLEED
45 PSI
RECIRC AIR OFF
APU
SYNOPTIC PAGE – AIR
CS1_CS3_2120_025
AIR PANEL
Figure 47: Recirculation System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
DETAILED DESCRIPTION The RFAN 115 VAC power is supplied from AC BUS 2 through a relay in electrical power center (EPC 2). The relay is controlled by control and distribution cabinet 2 (CDC 2). CDC 2 monitors the relay position using the discrete feedback from a relay contact. The position information is sent to the IASCs through the data concentrator unit module cabinet (DMC). The recirculation system operation is fully automatic. The RECIRC AIR PBA is used only in case of recirculation fan failure or in case of smoke detected in the avionics extraction system, and confirms shutdown of the recirculation system. The PBA commands the CDC 2 to isolate the 115 VAC power from the RFAN, bypassing IASC commands. Each IASC communicates with the RFAN controller through a CAN BUS to automatically adjust the fan speed. The IASC controls the RFAN speed according to a schedule as a function of system configuration •
Number of packs operating
•
Pull-up or pull-down sequence
•
Cabin altitude
•
Weight-on-wheels (WOW) signal
•
Cargo ventilation status
•
Aircraft version (CS100 or CS300)
If the FWD and AFT CARGO switches are selected OFF, with the aircraft operating on ground at a field elevation greater than 12,000 ft, the recirculation fan is automatically turned off by the IACSs to prevent excessive distribution noise. The normal maximum fan speed is 15,800 rpm. If the cabin altitude is out of its normal range, the RFAN speed increases to 16,000 rpm. The RFAN provides protection against internal overtemperature, overcurrent, and overvoltage. In case of a major internal failure, the RFAN is automatically turned off, and latched until the RFAN power is recycled. Once the RFAN fault is detected by the IASCs either from the controller or from the solid-state power controller (SSPC) feedback, an RECIRC AIR FAIL caution message is displayed.
NOTE 1. On the CS100, the recirculation fan is shutoff during the pull-up sequence. 2. On the CS300, the recirculation fan runs at maximum speed during the pull-up sequence.
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21 - Environmental Control 21-22 Recirculation
Recirc Fan AC BUS 2 20A RECIRCULATION FAN FIDEX CONTROL UNIT
RECIRC AIR
EPC 2 DMC 1 AND DMC 2
OFF
CDC 2-9-13 RECIRC FAN RLY CMD DC SSPC 3A BUS 2 Logic: Advanced
Recirc Fan Discrete Air/Gnd Recirc Fan Fault
CDC DMC EPC FIDEX IASC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electrical Power Center Fire Detection and Extinguishing Integrated Air Systems Controller AFDX ARINC 429 CAN BUS
Logic: Always On CDC 2 IASC 1 and IASC 2
CS1_CS3_2120_021
LEGEND
CDC 2-9-14 RECIRC FAN CTRL DC SSPC 3A BUS 2
Figure 48: Recirculation Fan Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Recirculation
MONITORING AND TEST The following page provides the crew alerting system (CAS) and INFO messages associated with the recirculation system:
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21 - Environmental Control 21-22 Recirculation
CAS MESSAGES Table 21: CAUTION Message MESSAGE
LOGIC
RECIRC AIR FAIL RFAN internal failure detected. RFAN digital bus communication failure detected. Loss of both IASC B channels with RFAN powered and IASC safety channels remain functional.
Table 22: STATUS Message MESSAGE
LOGIC
RECIRC AIR OFF Recirculation air selected OFF.
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21 - Environmental Control 21-22 Distribution System
21-22 DISTRIBUTION SYSTEM GENERAL DESCRIPTION The low-pressure distribution system consists of a network of rigid and flexible composite ducts, which collects air from the air supply sources, mixes fresh conditioned air with recirculation air in a mixing manifold unit. Air conditioning is normally supplied from two air conditioning packs through low-pressure distribution lines. Air conditioning can also be supplied from a low-pressure ground cart through the low-pressure ground connection. An emergency ram air valve connects the ram air inlet to the low-pressure distribution system just upstream the bulkhead check valves in flight. This provides ventilation to the cabin if both air conditioning packs fail. Two bulkhead check valves isolate the pressurized cabin from the bleed air and air conditioning packs installed in the unpressurized wing-to-body fairing (WTBF). In the event of a distribution line rupture outside of the pressurized compartment, the bulkhead check valves close to minimize pressure loss in the aircraft. Air from the air conditioning packs is mixed with recirculation air just before entering the mix manifold. The mix manifold has three main outlets which split the flow for distribution. The air is distributed to three zones; flight deck, forward cabin, and aft cabin through dedicated distribution lines.
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21 - Environmental Control 21-22 Distribution System
Low-Pressure Ground Connection
From Left Pack
Emergency Ram Air Valve
From Right Pack
Ram Air
Check Valves
To Cargo Ventilation MIX MANIFOLD
To Aft Cabin
To Flight Deck Air To Forward Cabin
From Recirculation Fan
CS1_CS3_2120_024
From Recirculation Fan
To Gasper Air
LEGEND Cold Air Recirculated Air
Figure 49: Distribution System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
ZONE DISTRIBUTION
Lavatory and Galley Areas
Both forward and aft cabin zones and the flight deck are supplied independently from the mixing manifold, allowing for three independent zone temperature controls in the aircraft.
The air distribution is provided by air outlets located in the galley and lavatory areas. The ventilation flow for the galleys and lavatories is provided at a lower rate than the extraction flow to prevent odors from the galleys and lavatories flowing toward the passenger seating area.
The air is ducted by risers into overhead stowage bins on both sides of the cabin, and then discharged into a plenum that distributes air at each cabin row from the top, the bottom, and the side of the bins. On each side of the cabin, an underfloor distribution plenum connects the risers for the flight deck, the forward cabin, and the aft cabin with the mixing manifold. The end of each plenum connects with the galley and lavatory areas located at both ends of the cabin. Conditioned air is distributed via a dedicated gasper air line. The line runs from the mix manifold to risers on the right and left sides of the aircraft. The risers connect to the overhead stowage bins for distribution to the gasper air outlets on the passenger service units.
Flight attendant gasper air outlets in the galley and entryway areas are connected to the distribution lines. There are provisions for three gasper outlets for the forward zone, and four gasper outlets for the aft zone. Flight Deck The flight deck air is distributed by a dedicated duct line that runs along the left side underfloor of the cabin and supplies the flight deck. For the CS100, the flight deck/cabin flow split is 14%/86% respectively.
CS100 Cabin Distribution There are 10 risers on each side for the forward cabin zone. The farthest forward riser on each side is dedicated to the forward galley area. The aft riser on each side provides extra flow for the overwing emergency exit (OWEE) doors area ventilation. There are 12 risers on each side for the aft cabin zone. The farthest aft riser on each side is dedicated to the aft galley area. CS100 Gasper Air The CS100 has five gasper risers. The three on the right and two on the left connect the mix manifold gasper air outlet to the overhead stowage bins.
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21 - Environmental Control 21-22 Distribution System
Risers
Flight Deck Distribution Fwd Cabin Distribution Aft Cabin Distribution Gasper Air CS100
CS1_CS3_2120_015
LEGEND
Figure 50: Zone Distribution CS100 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
CS300 Cabin Distribution There are 12 risers on each side for the farthest forward cabin zone. The forward riser on each side is dedicated to the forward galley area. The most aft riser on each side provides extra flow for the overwing emergency exit (OWEE) doors area ventilation. There are 12 risers on each side for the aft cabin zone. The farthest aft riser on each side is dedicated to the aft galley area. For the CS300, the flight deck/cabin flow split is set at 12.5%/87.5% respectively. CS300 Gasper Air The CS300 has seven gasper risers. The four on the right and three on the left connect the mix manifold gasper air outlet to the overhead stowage bins.
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21 - Environmental Control 21-22 Distribution System
Risers
Flight Deck Distribution Fwd Cabin Distribution Aft Cabin Distribution Gasper Air CS300
CS1_CS3_2120_027
LEGEND
Figure 51: Zone Distribution CS300 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
OVERHEAD BIN DISTRIBUTION The forward and aft cabin air is distributed from the top of the cabin through airflow channels integral to the passenger baggage overhead bins. The low-pressure distribution network connects to the bins by means of duct risers located of both side of the inner fuselage.
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21 - Environmental Control 21-22 Distribution System
50%
GASPER AIR FLOW
From riser 25%
25%
CABIN AIR DISTRIBUTION
CS1_CS3_2120_028
From riser
Figure 52: Overhead Bin Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
FLIGHT DECK DISTRIBUTION From the mix manifold, air is routed to the flight deck via underfloor low-pressure ducting, passing into the flight deck within a single duct to provide ventilation for the crew via gaspers as well as general ventilation via diffusers in the side panels. Each gasper outlet is adjustable for flow and direction. The ventilation air is also passed into the flight deck through the glareshield. The air supplied to the flight deck is ducted up through the floor at the center console, and routed to provide additional ventilation behind the control panels. In the flight deck the following air outlets are provided: •
Large upward facing outlet in each side console
•
Gasper outlets above the pilot, copilot, and observer head ventilation
•
Gasper outlet on each side of the forward instrument panel
•
Gasper outlet on each side of the floor pan at foot level
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21 - Environmental Control 21-22 Distribution System
Overhead Gasper Outlets
Side Console Ventilation
Side Console Ventilation Fwd Glareshield Outlets
Flight Deck Distribution Duct
Knee Gasper Outlet
CS1_CS3_2120_032
Side Gasper Outlet
Figure 53: Flight Deck Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
COMPONENT LOCATION The following distribution system components are located in the environmental control system compartment, aft of the forward cargo compartment. •
Mix manifold
•
Bulkhead check valve
•
Low-pressure ground connection
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21 - Environmental Control 21-22 Distribution System
Mix Manifold
Bulkhead Check Valve
CS1_CS3_2120_030
Low-Pressure Ground Connection
Bulkhead Check Valve
Figure 54: Distribution System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-125
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21 - Environmental Control 21-22 Distribution System
COMPONENT INFORMATION LOW-PRESSURE GROUND CONNECTION The low-pressure ground connection allows the connection of an external low-pressure ground cart to the aircraft during ground operations. The low-pressure ground cart supplies the aircraft cabin with conditioned air using the aircraft low-pressure distribution system. The low-pressure ground connection consists of an adapter, flapper check valve, and a cover. In flight, when the cabin is pressurized, the flapper check valve and adapter are designed to withstand differential pressure of 8.9 psid, which is slightly higher than the differential pressure seen by the pressurized vessel. A cover provides protection from water entering the main distribution system in the event of an aircraft ditching. The cover is held in place by a pin that remains attached to the aircraft by a lanyard.
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21 - Environmental Control 21-22 Distribution System
Check Valve
Pin Cover
CS1_CS3_2120_010
Adapter
Figure 55: Low-Pressure Ground Connection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-22 Distribution System
MIX MANIFOLD The mix manifold acts as a mixing chamber achieving, the required air temperature for distribution. Two air-conditioning packs feed fresh air into the mix manifold. This fresh air mixes with recirculated air collected from the underfloor area near the outflow valve. Some pre-mixing occurs prior to the mix manifold from the junction of the recirculation ducts and the pack ducts. The mix manifold has four main outlets, which split the flow for distribution to the forward cabin, aft cabin, flight deck, and gasper air.
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21 - Environmental Control 21-22 Distribution System
Mix Manifold Left Pack Air
Right Pack Air
Recirculation Air
Aft Cabin Distribution
Gasper Air
Flight Deck Distribution
CS1_CS3_2120_016
Fwd Cabin Distribution
Figure 56: Mix Manifold Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
21-23 VENTILATION SYSTEM GENERAL DESCRIPTION Air tapped off the recirculation air ventilates the cargo compartments through a cargo shutoff valve (CSOV). Each cargo compartment has two CSOVs. One is installed in the inlet duct, and the other is installed at the opposite end of the compartment in the exhaust duct. The CSOV is an electrically-actuated butterfly valve. Microswitches report the open and closed positions. The valve remains in its last position if no power is supplied. The airflow maintains the temperature above 4°C in the forward cargo, and 2°C in the aft cargo. The air is exhausted into the underfloor area surrounding the cargo compartments. The forward cargo compartment CSOV exhausts air close to the outflow valve to avoid recirculating odors when transporting live cargo. The valves are open for normal operation. When the recirculation fan is selected OFF, the integrated air system controllers (IASCs) close the CSOVs. If the forward cargo heat is selected, the forward CSOVs remain open, and air from the air conditioning units mix with cargo trim air. The CSOVs isolate the cargo compartment when smoke is detected and reported by the fire detection and extinguishing (FIDEX) control unit. When the CARGO FIRE PBA is pressed, the IASCs close the CSOVs for the selected cargo compartment.
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21 - Environmental Control 21-23 Ventilation System
AIR PANEL FWD
MAN TEMP
VENT
OFF
CARGO LO HEAT
AFT
OFF
VENT
CARGO FIRE PANEL
HI HEAT
ON
FWD
CARGO BTL
AFT
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
F I RE FIRE E
DMC 1 and DMC 2
IASC 2
AVAIL
F IR E FIRE
IASC 1
FIDEX CONTROL UINIT
FWD Cargo Upstream CSOV
LEGEND CSOV DMC IASC
Cargo Shutoff Valve Data Concentrator Unit Module Cabinet Integrated Air System Controller ARINC 429
FWD Cargo Downstream CSOV
Aft Cargo Downstream CSOV
CS1_CS3_2120_018
rgo Aft Cargo eam Upstream CSOV
Figure 57: Ventilation System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
COMPONENT LOCATION The following component is installed in the ventilation system: •
Cargo shutoff valves (CSOVs)
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21 - Environmental Control 21-23 Ventilation System
Upstream CSOV
Upstream CSOV
Fwd CSOV
CS1_CS3_2120_031
Downstream CSOV
Figure 58: Cargo Shutoff Valves Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
CONTROLS AND INDICATIONS AIR PANEL The forward and aft cargo compartment ventilation is controlled from the AIR panel. The forward CSOVs open anytime the FWD CARGO switch is selected to VENT, LO HEAT, or HI HEAT. The aft CSOVs open when the AFT CARGO switch is selected to VENT. CARGO PANEL If either FIRE PBA is pressed on the CARGO panel, the CSOVs are driven closed by the IASCs.
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21 - Environmental Control 21-23 Ventilation System
AIR COCKPIT
C
FWD
H
CABIN
C
AFT
H
C
FWD
MAN TEMP
CARGO
VENT
OFF
LO HEAT
H
AFT
OFF
VENT
HI HEAT
ON
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
FWD
F IRE IRE L PACK FAIL OFF
FAIL
BTL
AVAIL
AFT
F I RE RE
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
CARGO
FAIL OFF
CARGO PANEL
R BLEED FAIL
APU BLEED
OFF
FAIL CS1_CS3_2120_033
OFF
AIR PANEL
Figure 59: Ventilation System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
INDICATIONS Synoptic Page The FWD CARGO switch selection; OFF, VENT, LO, or HI is displayed on the AIR synoptic page. EICAS Page The forward cargo temperature indication on the engine indication and crew alerting system (EICAS) page is grayed out when the FWD CARGO switch is set to OFF or VENT.
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21 - Environmental Control 21-23 Ventilation System
STATUS FUEL
AIR HYD
DOOR
ELEC
AVIONIC
CKPT 23 °C 22 °C
VENT
15°C
10 °C
CARGO
CLB
FLT CTRL
INFO
FWD 22 °C 22 °C -+ 2
AFT 22 °C 22 °C -+ 2
15 °C
15 °C
CB
73.0
73.3
718
722
EGT
86.5 2410 124 116
R PACK
L PACK
92.2
N1
RAM AIR
TRIM AIR
92.2
86.6 2440 124 116
N2 FF (PPH) OIL TEMP OIL PRESS
TOTAL FUEL (LB) 10905 2735
45 PSI
XBLEED
5435
2735
45 PSI
1400 5.7 ELEV 570
CAB ALT
RATE
ó3
CREW
LDG TEMP (°C)
23
0 OXY 1850
TRIM NU STAB
22
22
ND
4.2 NL RUDDER NR
APU
SYNOPTIC PAGE – AIR
EICAS
CS1_CS3_2123_001
INFO
Figure 60: Ventilation System Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
DETAILED DESCRIPTION The cargo shutoff valves are tested at landing by the IASCs, which perform a full open-close cycle to check the motor operation as well as the full close and full open position feedback. Only the in-control channel driver of each IASC is tested during the test. If a failure is detected, switching to the other channel is performed, and the CSOV cycling is commanded again. During normal operation, if the CSOV does not respond to the IASC command, an INFO message is displayed
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
AIR COCKPIT
C
FWD
H
C
CABIN
H
C
FWD
MAN TEMP OFF
VENT
AFT
CARGO LO HEAT
FIDEX CONTROL UNIT
H
FWD
AFT
OFF
PACK FLOW
RECIRC AIR
OFF
HI
OFF
AVAIL
FII R E F
AFT
FII R E F
DMC 1 and DMC 2
IASC 1 TRIM AIR
BTL
VENT
HI HEAT
ON
CARGO
CARGO PANEL OPEN
Position
IASC 2
AIR PANEL
RECIRCULATION FAN
MIX MANIFOLD
Downstream CSOV
Downstream CSOV AFT CARGO COMPARTMENT
LEGEND CSOV DMC FIDEX IASC
From Right Pack
Cargo Shutoff Valve Data Concentrator Unit Module Cabinet Fire Detection and Extinguishing Integrated Air System Controller ARINC 429
Power
FWD CARGO COMPARTMENT
Upstream CSOV
Upstream CSOV
From Trim Air System
CS1_CS3_2120_020
From Left Pack
Figure 61: Ventilation System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-23 Ventilation System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the ventilation system:
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21 - Environmental Control 21-23 Ventilation System
CAS MESSAGES Table 23: STATUS Messages MESSAGE
LOGIC
FWD CARGO AIR FWD cargo heating selection selected and confirmed OFF OFF. AFT CARGO AIR OFF
AFT CARGO selected and confirmed OFF.
Table 24: INFO Messages MESSAGE
LOGIC
21 AIR SYSTEM FAULT - FWD CARGO SOV INOP
Upstream of downstream forward cargo SOV failed in position or in unknown position.
21 AIR SYSTEM FAULT - AFT CARGO SOV INOP
Upstream of downstream forward cargo SOV failed in position or in unknown position.
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21 - Environmental Control 21-23 Ventilation System
PRACTICAL ASPECTS The cargo shutoff valve (CSOV) has a manual lever used to secure in the closed position before dispatch if the valve has failed.
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21 - Environmental Control 21-23 Ventilation System
CS1_CS3_2123_003
Manual Shutoff Lever
Figure 62: Cargo Shutoff Valve Practical Aspects Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
21-40 HEATING FLIGHT CREW FLOOR HEATING GENERAL DESCRIPTION The flight crew floor heating system consists of heated mats embedded in each foot rest plate. The heated mats operate as soon as the main DC buses are powered. Thermostats embedded in the mats control the operation of the mats.
COMPONENT LOCATION •
Heated mat
HEATED MAT The heated mats are installed in the foot rest plates in front of each pilot seat.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control
Pilot Left Heated Mat
Pilot Right Heated Mat
Copilot Left Heated Mat
Copilot Right Heated Mat
CS1_CS3_2140_004
21-40 Heating
Figure 63: Flight Crew Floor Heating General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
DETAILED DESCRIPTION The heated mats are controlled by thermostats that provide power when the temperature is less than 20°C (68°F) and remove power when the temperature exceeds 40°C (104°F). The control and distribution cabinet 3 (CDC 3) and CDC 4 monitor the current draw. When the current draw exceeds 1 A, the solid-state power controller (SSPC) latches off and reports the overcurrent fault to the onboard maintenance system (OMS).
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21 - Environmental Control 21-40 Heating
CDC 3 CDC 3-4-4 L FOOT WARMER DC SSPC 3A BUS 1
HEATED MAT 1 Pilot
Logic: Advanced HEATED MAT 2
CDC 4 CDC 4-4-4 R FOOT WARMER DC SSPC 3A BUS 2
HEATED MAT 3 Co-pilot
Logic: Advanced HEATED MAT 4 CDC 1 and CDC 2
IASC 1 and IASC 2
DMC 1 and DMC 2
OMS
CDC DMC IASC OMS
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Onboard Maintenance System AFDX ARINC 429 TTP BUS
CS1_CS3_2140_005
LEGEND
Figure 64: Flight Crew Floor Heating Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
GALLEY HEATING GENERAL DESCRIPTION The forward and aft galley heating consists of a fan and a heater in a recirculation loop that distributes reheated air next to the service and passenger doors. Air is extracted from the galley area by the galley fan, heated by the galley heater and then exhausted near the floor next to the service and main passenger doors.
Each forward or aft galley heating system is controlled from the flight attendant panels located in the respective forward and aft galley areas. The system status is sent from IASC to the cabin management system (CMS) for fault indication.
The respective IASC monitors the supplemental heating system using a dual-element galley heater temperature sensor (GHTS). The sensor limits the duct temperature to 70°C. Three modes of operation can be selected on a panel installed in the galley area: •
When FAN ON is selected, the fan operates
•
When LO HEAT is selected, the fan operates and the heater operates on a two phase power. With a cabin inlet temperature of 25°C (77°F), the system provides an outlet temperature between 30°C (86°F) and 40°C (104°F)
•
When HI HEAT is selected, the fan operates and the heater operates on three-phase power. With a cabin inlet temperature of 25°C (77°F), the system provides an outlet temperature between 45°C (113°F) and 55°C (131°F)
A dual-temperature sensor is installed at the heater outlet to provide the duct temperature to the IASCs for system monitoring. The forward galley heater receives 3-phase 115 VAC from AC BUS 1 and the aft galley heater receives 3-phase power from AC BUS 2. The forward galley fan is powered by DC BUS 1 and the aft galley fan is powered by DC BUS 2.
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21 - Environmental Control 21-40 Heating
FWD GALLEY PANEL FAN ON
HEATER
IASC 1 CDC 3
HIGH LOW
AC BUS 1 DC BUS 1
ZONE BOX 1
FAN FWD ENTRY WAY
CABIN CONTROLLER
FWD GALLEY
DMC 1 AND DMC 2
HEATER
GALLEY HEAT TEMPERATURE SENSOR CDC 5 AC BUS 2 DC BUS 2
ZONE BOX 2 IASC 2
ON
LEGEND CDC DMC IASC
HEATER HIGH
FAN
LOW
AFT ENTRY WAY
AFT GALLEY PANEL
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 Ethernet
AFT GALLEY
HEATER
GALLEY HEAT TEMPERATURE SENSOR
CS1_CS3_2140_002
FAN
Figure 65: Galley Heating Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
COMPONENT LOCATION The galley heating system consists of: •
Galley fan
•
Galley heater
•
Galley heater temperature sensor
GALLEY FAN The galley fan is attached to the galley structure. GALLEY HEATER The galley heater is attached to the galley structure. GALLEY HEATER TEMPERATURE SENSOR The galley temperature sensor is installed in the duct at outlet of the galley heater.
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21 - Environmental Control 21-40 Heating
Galley Fan
Heater
Temperature Sensor NOTE Fwd galley shown. Aft galley similar.
Outlet
CS1_CS3_2120_014
Outlet
Figure 66: Supplemental Galley Heat Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
CONTROLS AND INDICATIONS GALLEY HEAT CONTROLS The galley heaters are controlled from the GALLEY panels located in each galley. Each heater has a fan control and a HIGH or LOW heat selection.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
GA GALLEY ALLEY SHUTOFF SHU UTO OFF LIGHTS
BRT
GALLEY COUNTER
DIM BRT
GALLEY AREA
DIM
FAN
HEATER
HIGH
ON
LOW CHILLER
DEFROST FAIL CS1_CS3_2140_007
ON
Figure 67: Galley Heater Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
GALLEY HEAT INDICATIONS The galley heater faults are displayed on the CMS screen.
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21 - Environmental Control
CS1_CS3_2140_008
21-40 Heating
Figure 68: Galley Heater Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
DETAILED DESCRIPTION The heater operation is regulated by a built-in thermostat. For overtemperature protection, thermal fuses in-line with the heating elements open at 183°C (362°F). A heater fault discrete output from the heater assembly is connected directly to the control and distribution cabinet (CDC) that is its power. The galley heater 115 VAC power is removed when a fault occurs. The galley fan remains in operation and only ventilation is provided by the system. A fan fault discrete output signals overcurrent, overvoltage, underspeed (10,000 rpm) or overheat conditions in the fan assembly, which is connected directly to the CDC that provides power to the galley fan. Both the 28 VDC power to the galley fan and 115 VAC power to the galley heater are removed when a fault occurs. In both cases, the fault data is sent to the onboard maintenance system (OMS), and a GALLEY HEATING SYSTEM UNAVAILABLE message is displayed on the cabin management system (CMS) crew terminal. The IASC monitors the galley heater temperature. The galley heater temperature sensor has two elements. One element reports to each of the IASC channels. In case of overheat conditions detected by either element of the galley heater temperature sensor, the heater power is turned off. The IASC commands the CDC to shut off the galley heating when: •
Galley heater outlet temperature reads above 70°C for more than 1 minute
•
The overheat condition is latched until the temperature decreases below 70°C and the GALLEY switch is cycled.
If both elements of the sensor fail, or one element and the opposite IASC channel fail, the heater power is removed. When both elements of the GHTS have failed, the heater is latched off until the GHTS is replaced.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-40 Heating
CDC 3 CDC 3-12-4,5,6 FWD GLY HTR AC SSPC 5A BUS 1
3Ø
Logic: Advanced
FWD GALLEY HEATER
Fwd Galley Heater Fault
FWD GALLEY PANEL FAN
HEATER
CDC 3-6-1 FWD GLY FAN DC BUS 1 SSPC 15A Logic: Advanced
HIGH
ON
FWD GALLEY FAN
Fwd Galley Fan Fault
LOW
ZONE BOX 1
CDC 1
CABIN CONTROLLER
DMC 1 and DMC 2
CDC DMC IASC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 Ethernet TTP
IASC 1 Channel A Channel B
FWD GALLEY HEATER TEMPERATURE SENSOR
NOTE Fwd galley shown. Aft galley similar.
CS1_CS3_2140_003
LEGEND
Figure 69: Supplemental Galley Heat Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
21-26 AVIONICS COOLING AND EXTRACTION SYSTEM GENERAL DESCRIPTION The avionics cooling and extraction system consists of three subsystems: •
Air inlet cooling system
•
Air extraction system
•
Backup exhaust and extraction system
In addition, the wing-to-body fairing (WTBF) is ventilated using the air conditioning ram air system. The air inlet cooling system draws air from the galleys, and underfloor area, cools it, and supplies it to the forward and mid avionics bay. The air extraction system draws air from behind the flight deck displays, both forward and mid equipment bays, control and distribution cabinet 5 (CDC 5) and the forward and aft lavatories. The backup exhaust and extraction system provides an alternate means to extract air from behind the flight deck displays, and both forward and mid equipment bays.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
OUTFLOW VALVE
FWD GALLEY
UNDERFLOOR
UNDERFLOOR
AFT GALLEY
CDC 5
M GROUND VALVE AVIONICS FILTER
AVIONICS FILTER
FWD SKIN HEAT EXCHANGER
EXTRACTION FAN 1
EXTRACTION FAN 2
MID BAY FAN
FWD BAY FAN
FWD EQUIPMENT BAY
AVIONICS COALESCENT FILTER
MID SKIN HEAT EXCHANGER
Flight Deck Displays
MID EQUIPMENT BAY
O2 Bottle Compartment
CDC 5
FWD LAV
AFT LAV
FLIGHT DECK
M
FWD AVIONICS EXHAUST VALVE
MID AVIONICS EXHAUST VALVE
M
BACKUP FAN
LEGEND Air Inlet Cooling System Backup Exhaust and Extraction System Air Extraction System
CS1_CS3_2126_012
UNDERFLOOR
Figure 70: Avionics Cooling and Extraction System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
AIR INLET COOLING SYSTEM GENERAL DESCRIPTION The air inlet cooling system supplies cooled and filtered air into both forward and mid equipment bays in order to lower their ambient operating temperature. There is one air inlet cooling system dedicated to each equipment bay. Each system has a fan, skin heat exchanger, and a filter. The fan extracts air from the galley and underfloor areas. The air is cleaned by a filter before passing through the skin heat exchanger. The skin heat exchanger cools the air using the cold aircraft skin. The cooled air is distributed through piccolo tubes in the equipment bay to cool the equipment. The forward bay fan is powered by DC BUS 2. The mid bay fan is powered by AC BUS 2. The mid bay air inlet cooling system has a ground valve that provides additional air to the mid equipment bay when the air inlet cooling system operates on the ground. The ground valve is powered by DC BUS 2. On the ground, approximately 70% of the cooling comes from the ground valve, and the remaining 30% comes from the galley and control and distribution cabinet (CDC) 5. A coalescent filter is installed downstream of a ground valve, located on the aircraft skin. The coalescent filter extracts the water contained in the outside air and then filters the air for the mid equipment bay. Control of the system is fully automatic. The INLET PBA is normally left in the auto position and only selected OFF when the system has automatically reconfigured to the OFF position. Selecting the INLET PBA OFF turns off the forward and mid bay fans, and closes the ground valve.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
DMC 1 and DMC 2
IASC 1
CDC 4 CMD/FAULT UNDERFLOOR
IASC 2
CDC 2 AC BUS 2
FWD GALLEY
CDC 5
AFT GALLEY M GROUND VALVE
AVIONICS FILTER
FWD SKIN HEAT EXCHANGER
MID SKIN HEAT EXCHANGER
AVIONICS COALESCENT FILTER
MID BAY FAN
FWD BAY FAN DC BUS 2 FWD EQUIPMENT BAY
MID EQUIPMENT BAY
LEGEND
CDC DMC IASC
Inlet Air Cooling Air Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 Discrete TTP
CS1_CS3_2126_014
AVIONICS FILTER
Figure 71: Air Inlet Cooling Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
COMPONENT LOCATION
MID BAY FAN
The following are major components of the air inlet cooling system:
The mid bay fan is located in the aft cargo compartment on the right side of the fuselage, forward of the cargo door.
•
Forward avionics filter
•
Forward bay fan
AVIONICS COALESCENT FILTER
•
Forward skin heat exchanger
•
Ground valve
The avionics coalescent filter is located in the aft cargo compartment on the right side of the fuselage, forward of the cargo door.
•
Mid bay fan
MID SKIN HEAT EXCHANGER
•
Avionics coalescent filter
•
Mid skin heat exchanger
The aft skin heat exchanger is located in the aft cargo compartment on the right side of the fuselage, aft of the cargo door.
•
Mid avionics filter
MID AVIONICS FILTER
•
Ducting
The mid avionics filter is located in the aft cargo compartment on the right side of the fuselage aft, of the cargo door.
FORWARD AVIONICS FILTER The forward avionics filter is located on the right side of the forward equipment bay. FORWARD BAY FAN The forward bay fan is located on the right side of the forward equipment bay. FORWARD SKIN HEAT EXCHANGER
DUCTING Two separate duct networks provide air to each equipment bay. Each duct runs along the left side of the fuselage. The forward duct connects the forward galleys and underfloor area to the forward equipment bay. The mid duct connects the aft galley and control and distribution cabinet (CDC 5) to the mid equipment bay. The mid bay also has a duct that draws external air when the aircraft is on the ground. The air is distributed in the equipment bays through piccolo tubes.
The forward skin heat exchanger is located on the right side of the forward equipment bay. GROUND VALVE The ground valve is located in the aft cargo compartment on the right side of the fuselage, forward of the cargo door.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Galley 4
Mid Skin Heat Exchanger
Ground Valve
Avionics Coalescent Filter Galley 4A (if installed) CDC 5
Mid Bay Fan
Mid Avionics Filter
Ducts
Mid Equipment Bay Piccolo Tubes
Galley 2 Galley 1
Fwd Skin Heat Exchanger
Fwd Bay Fan
CS1_CS3_2126_013
Fwd Avionics Filter
Fwd Equipment Bay Piccolo Tubes
Figure 72: Air Inlet Cooling Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
COMPONENT INFORMATION FORWARD AVIONICS FILTER The forward avionics filter filters air from the forward galleys and underfloor area. The replaceable filter is installed at the inlet to the forward skin heat exchanger. FORWARD SKIN HEAT EXCHANGER The forward skin heat exchanger is installed within the frames of the fuselage structure.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Fwd Avionics Filter
CS1_CS3_2126_005
Fwd Skin Heat Exchanger
Figure 73: Forward Avionics Filter and Forward Skin Heat Exchanger Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
GROUND VALVE The GND VLV is installed on the aircraft skin on the lower right side of the fuselage. The ground valve is driven by a linear electrical actuator powered by 28 VDC. A torque limiter protects the actuator motor and prevents inadvertent opening of the valve in flight. The valve position is monitored by microswitches. The valve has a manual override handle on its external side. This handle allows the valve to be closed from outside of the aircraft. AVIONICS COALESCENT FILTER The avionics coalescent filter is located in the duct from the ground valve. The replaceable filter removes contaminants and water from the outside air. A drain line carries the water to the lower fuselage to drain overboard.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Avionics Coalescent Filter GROUND VALVE
Manual Close Lever
CS1_CS3_2126_006
Filter Drain
Figure 74: Ground Valve and Avionics Coalescent Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
MID SKIN HEAT EXCHANGER The mid skin heat exchanger is installed within the frames of the fuselage structure. MID AVIONICS FILTER The mid avionics filter filters air drawn from the aft galley and underfloor area. The replaceable filter is installed at the inlet to the mid skin heat exchanger.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
Mid Avionics Filter
CS1_CS3_2126_007
Mid Skin Heat Exchanger
Figure 75: Mid Skin Heat Exchanger and Mid Avionics Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
CONTROLS AND INDICATIONS The system is controlled from the EQUIP COOLING panel in the flight deck. In normal operation, the system operation is automatically controlled by the integrated air system controllers (IASCs). The INLET PBA can be selected to turn off the air inlet cooling system.
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
CARGO
FWD
BTL
AVAIL
FIRE
AFT
FIRE
OVERHEAD PANEL
EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
PRESSURIZATION EMER DEPRESS
CS1_CS3_2126_011
ON
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Figure 76: Avionics Cooling and Extraction System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-26 Avionics Cooling and Extraction System
DETAILED DESCRIPTION The air inlet cooling system operates all the time on ground and in flight. The air inlet cooling system operation is automatically managed by the IASC main channels, which control the mid ground valve and both forward and mid bay inlet fans through logics embedded in the CDCs. The CDCs drive power supplies, discrete and relay signals located in the electrical distribution system according to requests from the IASCs.
In case a failure is detected on either forward or mid bay fan or the ground valve, an advisory message EQUIP BAY COOL FAULT is displayed. An INFO message provides guidance for any action to be taken before the next aircraft dispatch. The air inlet cooling system can be selected OFF by selecting the INLET PBA on the overhead panel. The aircraft can be dispatched without any operating limitation with the system selected OFF.
In flight, the air inlet cooling system extracts air from the galley, underfloor areas, and CDC 5, cools it through the skin heat exchanger, and then supplies it through a set of distribution lines to cool the avionics bays. On ground, with an outside air temperature (OAT) ranging between 2°C and 33°C, the IASCs command the ground valve open to allow outside air into the system. Before takeoff, the ground valve is driven closed once all doors are signaled closed and locked. If the ground valve fails to close, the valve can be manually closed from outside of the aircraft. After landing, the ground valve is driven open once all doors are not signaled closed. The ground valve is kept closed at landing until the doors open in order to minimize the residual cabin differential pressure, and prevent engine exhaust air going into the filter during the deceleration and taxi phases. In case of smoke detected by the fire detection and extinguishing system (FIDEX) control unit, the IASCs signal the CDCs to remove power from both forward and mid bay fans to automatically shutoff the air inlet cooling system. In case of ditching, when the DITCHING PBA is selected on the PRESSURIZATION panel, the ground valve is automatically driven closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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FWD BAY INLET FAN ON
On/Off
Fault CDC 4
EQUIP COOLING
FWD BAY IN FAN DC BUS 2
EXHAUST VLV AUTO ON ONLY
CDC 2-17-1, 2, 3 MID BAY in AUX AC BUS 2 SSPC 10A
FIDEX CONTROL UNIT
Logic: Always On
PRESSURIZATION EMER DEPRESS
CDC 2-9-15 MID INLET VLV DC SSPC 3A BUS 2
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Logic: Advanced Fault
DITCHING FAIL MAN
LEGEND ON
CDC DMC EPC FIDEX IASC
115 VAC
Logic: Advanced Fault CDC 2-9-9 MID BAY in FAN CMD DC SSPC 3A BUS 2
DMC 1 and DMC 2
AUTO PRESS
28 VDC FWD EQUIPMENT BAY INLET FAN
EPC 2
OFF
ON
25
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electrical Power Center Fire Detection and Extinguishing Integrated Air System Controller AFDX ARINC 429 TTP BUS
MID EQUIPMENT BAY INLET FAN
28 VDC
CDC 2
Position IASC 2
GROUND VALVE
CS1_CS3_2126_019
INLET
Figure 77: Air Inlet Cooling System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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AIR EXTRACTION SYSTEM GENERAL DESCRIPTION The extraction system is composed of two identical extraction fans (EFANs) pulling air from the flight deck displays, both forward and mid equipment bays, CDC 5 enclosure, and the forward and aft lavatories. Extracted air is discharged in closed vicinity of the outflow valve, and then dumped overboard. The extracted airflow is kept below 70°C. The air extraction system operates when the EXHAUST switch on the EQUIP COOLING panel is in the AUTO position. In normal operation, one fan is running, the other is on standby-mode. For reliability and safety reasons, the fans are alternately operated on daily basis. EFAN 1 is powered by AC BUS 2. The motor controller is powered by DC BUS 2. EFAN 2 is powered by AC BUS 1. The motor controller is powered by DC BUS 1. On the ground, the EFAN operates at maximum speed to maximize heat extraction, and maintain the equipment bay temperature as low as possible. In flight, the EFAN operates at a lower speed to maintain extraction flow below the outflow valve (OFV) exhaust flow value.
NOTE On the ground, powering the aircraft with batteries only must be limited to 5 minutes. After 5 minutes, AC power must be provided for EFAN operation to ensure adequate equipment bay cooling.
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EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
Safety CH
DMC 1 and DMC 2
IASC 1 CH A
Safety CH IASC 2 CH A
FLIGHT DECK Flight Deck Displays
O2 Bottle Compartment
FWD LAV
AFT LAV
FWD EQUIPMENT BAY
MID EQUIPMENT BAY
DC BUS 2
CDC 5
EXTRACTION FAN 1
AC BUS 2
CDC DMC IASC
CDC 2
Exhaust Air DC BUS 1 Control and Distribution Cabinet Data Concentrator Unit Module Cabinet AC BUS 1 Integrated Air System Controller ARINC 429 CDC 1 Discrete
OUTFLOW VALVE
EXTRACTION FAN 2
CS1_CS3_2126_016
LEGEND
Figure 78: Air Extraction System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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COMPONENT LOCATION The following is a major component of the main air extraction system •
Extraction fans
•
Ducting
EXTRACTION FANS The extraction fans are located in the environmental control system bay. DUCTING The extraction ducting runs along both side of the fuselage, connecting the flight deck display lines, oxygen compartment, and lavatories. Piccolo tubes draw air from the forward and mid equipment bays. The ducts join in a common manifold in the environmental equipment bay. The manifold connects all of the ducting to the extraction fans. The extraction fan exhaust discharges the exhaust air into the environmental bay near the outflow valve.
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Control and Distribution Cabinet 5
Mid Equipment Bay Piccolo Tubes
Lavatory E Extraction Fans
Fwd Equipment Bay Piccolo Tubes Flight Displays and Center Console
Extraction Fan Exhaust
CS1_CS3_2120_004
Ducts
Lavatory A Oxygen Compartment
Figure 79: Main Air Extraction System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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DETAILED DESCRIPTION In normal operation, one extraction fan (EFAN) is running, and the other extraction fan is on standby mode. Both avionics exhaust valves (AEV) are closed and the backup fan is off. The EFAN selection and the system reconfiguration are controlled by the IASC safety channels. Each safety channel controls the power supply of one EFAN by commanding the relay coil that switches the 115 VAC to the EFAN. Each EFAN is operated on different days. Automatic switching occurs on the ground. The fan selection logic is done by the IASCs safety channel, based on the date provided by the data concentrator unit module cabinet (DMC). Each extraction fan provides low-speed and internal fault monitoring feedback to the IASCs through a CAN BUS for troubleshooting and monitoring of the EFANs. A fault discrete signal is also sent from the EFAN to the IASC safety channels for monitoring of the EFAN internal failures. Fault detection is done by the fan motor control. Fan speed, overheat, overcurrent, and overvoltage faults are monitored. If one EFAN fails, the system automatically switches to the other EFAN, and an EQUIP BAY COOL FAULT advisory message is displayed. On the ground, the EFAN operates at full speed in order to maximize heat extraction, and maintain equipment bay ambient temperature as low as possible. In flight, the EFAN operates at a lower speed to maintain extraction flow below the outflow valve exhaust flow, and to ensure no equipment bay airflow is recirculated into the distribution system. Dispatch is possible with one EFAN failure, if the backup exhaust system is operative.
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EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
DMC 1 and DMC 2
IASC 1 EFAN 1 Fault
EFAN 1
EXTR Fan CTRL DC BUS 2
5
20 EXTR Fan 1 AC BUS 2
20 20
External FAN 1 Discrete CDC 1
NOTE EFAN 1 shown. EFAN 2 similar.
EPC 2
LEGEND Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Extraction Fan Electrical Power Center Integrated Air System Controller AFDX ARINC 429 CAN BUS
CS1_CS3_2126_017
CDC DMC EFAN EPC IASC
Figure 80: Main Air Extraction System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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BACKUP EXHAUST AND EXTRACTION SYSTEM GENERAL DESCRIPTION A backup cooling system is provided for the flight deck displays and avionics cooling functions with the addition of two duct networks, extracting air from behind the flight deck displays, and from the forward equipment bay and the mid equipment bay. Each duct connects to one avionics exhaust valve (AEV) installed on the aircraft skin. The AEV is a shutoff valve, which opens and exhausts the airflow overboard. Each AEV is powered and monitored by both IASCs. The amount of airflow extracted is mainly dependent on the cabin to ambient pressure differential existing between the inlet and the outlet of a flow restrictor integral to the aircraft skin. A third backup extraction fan is fitted on top of the mid avionics bay enclosure, and provides air extraction through an opening in the center ceiling of the mid avionics bay. The backup fan is powered by DC ESS BUS 1.
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EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
OFF
DMC 1 and DMC 2
IASC 2 CH A
CH B
IASC 1 CH A
CH B
FLIGHT DECK O2 Bottle Compartment
DC ESS BUS 1 BACKUP FAN
VENTS
MID EQUIPMENT BAY
FWD EQUIPMENT BAY
M
FWD AVIONICS EXHAUST VALVE
VENTS
M
MID AVIONICS EXHAUST VALVE
LEGEND DMC IASC VENTS
Exhaust Air Data Concentrator Unit Module Cabinet Integrated Air System Controller Ventilated Temperature Sensor ARINC 429 Discrete
CS1_CS3_2126_015
Flight Deck Displays
Figure 81: Backup Exhaust and Air Extraction System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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COMPONENT LOCATION The following are major components of the backup exhaust and air extraction system: •
Avionics exhaust valves
•
Backup fan
•
Ducting
•
Avionics bay ventilated temperature sensors (Refer to figure 83)
AVIONICS EXHAUST VALVES The forward avionics exhaust valve is located in the forward cargo compartment on the right side of the lower fuselage, forward of the cargo door. The aft avionics exhaust valve is located in the aft cargo compartment on the right side of the lower fuselage, forward of the cargo door. BACKUP FAN The backup fan is mounted on the ceiling of the mid equipment bay. DUCTING There are two duct networks for the backup exhaust and air extraction system. The forward duct runs along the left side of the aircraft, connecting the flight deck displays, and forward equipment bay to the forward avionics exhaust valve. The mid equipment bay ducting connects to the mid avionics exhaust valve. Air from the flight deck displays, forward, and mid equipment bays is drawn through piccolo tubes, and exhausted overboard when the avionics exhaust valves are open.
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Mid Avionics Exhaust Valve Backup Fan
Fwd Avionics Exhaust Valve Ducts Flight Deck Display Piccolo Tubes
CS1_CS3_2120_005
Ducts
Mid Equipment Bay Piccolo Tubes
Fwd Equipment Bay Piccolo Tubes
Figure 82: Backup Exhaust and Air Extraction System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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AVIONICS BAY VENTILATED TEMPERATURE SENSORS An avionics ventilated temperature sensor (AV-VENTS) is installed in each of the forward and mid equipment bay ceilings.
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Note
Fwd equipment bay
Fwd equipment bay shown. Mid equipment bay similar.
CS1_CS3_2126_010
Ventilated Temperature Sensor
Figure 83: Avionics Bay Ventilated Temperature Sensor Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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DETAILED DESCRIPTION The backup exhaust and extraction systems are in standby during normal operation. Each bay has an avionics ventilated temperature sensor (AV-VENTS) sensor for monitoring. The AV-VENTS sensor has two temperature elements with one reporting to each IASC. A small fan draws air across the elements. If the temperature exceeds 67°C (153°F) for more than 30 seconds, an EQUIP BAY OVHT warning message is displayed, and the auxiliary power unit (APU) (multifunction) horn sounds. The message is cleared when the temperature decreases to 62°C (144°F), and the exhaust switch on the EQUIP COOLING panel is moved from the AUTO position.
If the DITCHING PBA is selected on the PRESSURIZATION panel, the AEVs are automatically driven closed to prevent water from entering the aircraft. Both AEVs and the backup fan are tested 3 minutes after landing to confirm the backup exhaust and extraction systems is operative. In case a failure is detected on either AEV or on the backup fan at landing, an EICAS EQUIP BAY COOL FAULT advisory and relevant INFO messages are displayed. One or both AEVs can be failed for dispatch. Aircraft dispatch is not possible with the backup fan inoperative.
If both EFANs fail, or both channels of one AV-VENTS fail, the IASCs automatically open both avionics exhaust valves (AEVs), starts the backup fan, and an EICAS EQUIP BAY COOL FAIL caution message is displayed. The EXHAUST switch on the EQUIP COOLING panel is set to ON to match the new system configuration. In the event of smoke detected in the avionics bay, both AEVs are automatically driven open and the backup fan is kept off by the IASCs if no overheat is detected. The EXHAUST switch is moved to VLV ONLY to match the system configuration. A hardwired contact on the EXHAUST switch provides additional control of the fan.
NOTE When the EXHAUST switch is set to ON, a discrete hardwired signal from the switch turns on the backup fan. On ground when operating on the batteries only, as well as during approach when the ram air turbine (RAT) generator switches off, the electrical distribution system disables the backup fan in order to protect the batteries.
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ECS Safety Sensor DC ESS BUS 3
LEGEND CTL DMC EPC FIDEX IASC TEMP TLC VENTS
Control Data Concentrator Unit Module Cabinet Electrical Power Center Fire Detection and Extinguishing Integrated Air System Controller Temperature TRU Line Contactor Ventilated Temperature Sensor ARINC 429
3
RIGHT CIRCUIT BREAKER PANEL
Fwd Temp Mid Temp
EQUIP COOLING INLET
EXHAUST VLV AUTO ON ONLY
Backup Fan Ctl
Cmd
FWD AVIONICS EXHAUST VALVE
IASC 1
FIDEX CONTROL UNIT
DMC 1 and DMC 2
Open/Close Position
OFF
Mid Bay Altn Fan DC ESS BUS 1
FWD EQUIPMENT BAY VENTS
Fault 20
28 VDC
EPC 1 TLC 3 Aux Fan Ctl In TLC 3 Aux Fan Ctl Out EPC 3
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Backup Fan Ctl
DITCHING
MAN
Fwd Temp
Cmd
AUTO PRESS FAIL
Mid Temp MID EQUIPMENT BAY VENTS Open/Close Position
ON
MID AVIONICS EXHAUST VALVE IASC 2
CS1_CS3_2126_018
PRESSURIZATION
Fan on/off BACKUP FAN
Figure 84: Backup Exhaust and Air Extraction System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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COMPONENT INFORMATION AVIONICS EXHAUST VALVES The avionics exhaust valve (AEV) is an electrically actuated butterfly valve powered by a 28 VDC motor. The valve remains in its last position if power is lost. Microswitches provide position feedback to the IASCs. The AEV has a manual lever which can used to secure the valve in the open position after a failure.
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Manual Lever
Adapter
CS1_CS3_2126_009
Screen
Figure 85: Avionics Exhaust Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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BACKUP FAN The backup fan is mounted on the ceiling of the mid equipment bay. The fan runs at a single-speed when powered. A check valve closes off the fan exhaust when the fan is not running. The backup fan deflector is installed at the backup fan exit to direct smoke away from the floor sills where it could enter the passenger cabin.
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Check Valve
Mid Equipment Bay Ceiling
Backup Fan Deflector
CS1_CS3_2126_008
Backup Fan
Figure 86: Backup Fan Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CONTROLS AND INDICATIONS The system is controlled from the EQUIP COOLING panel in the flight deck. In normal operation, the EXHAUST rotary switch is in the AUTO position and the system operation is automatically controlled by the IASCs. If a fault occurs in the main extraction system, the IASC automatically selects the backup exhaust and extraction system. The EXHAUST rotary switch is moved to the ON position to match the system configuration. If smoke is detected in an avionics bay, the backup extraction fan can be selected off using the VLV ONLY rotary switch position.
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CARGO
FWD
BTL
AVAIL
FIRE
AFT
FIRE
EQUIP COOLING OVERHEAD
INLET
EXHAUST VLV AUTO ON ONLY
OFF
PRESSURIZATION EMER DEPRESS
CS1_CS3_2126_020
ON
MAX ∆P 0.11 PSI TO 1.00 PSI LDG
Figure 87: Avionics Cooling and Extraction System - Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the avionics cooling and extraction system:
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CAS MESSAGES
Table 28: STATUS Messages Table 25: WARNING Message
MESSAGE EQUIP BAY OVHT
MESSAGE INLET AIR OFF
LOGIC Fwd or mid avionics bay overtemperature above 67° C.
EQUIP BAY COOL FAIL
MESSAGE
LOGIC Loss of both extraction FANS confirmed by either IASC control or safety channel. Both FANS are running. Total loss of one zone temperature measurement, with electrical power supplied, and with either IASC safety channel operational.
Table 27: ADVISORY Messages MESSAGE EQUIP BAY COOL FAULT
LOGIC FWD or MID equipment bay cooling fault
Table 28: STATUS Messages MESSAGE
LOGIC
EXHAUST AIR ON
Exhaust switch selected to ON. Supplemental fan turns on and valves open.
EXHAUST AIR VLV ONLY
EXHAUST selector to VLV ONLY and both AEV commanded OPEN.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
INLET PBA selected OFF and both forward and mid inlet fans commanded OFF and the mid ground valve commanded closed.
Table 29: INFO Messages
Table 26: CAUTION Message MESSAGE
LOGIC
LOGIC
21 EQUIP BAY COOL FAULT EFAN INOP
Extraction FAN 1 or FAN 2 inoperative.
21 EQUIP BAY COOL FAULT EFAN CAN BUS INOP
Extraction FAN 1 or FAN 2 CAN BUS inoperative.
21 EQUIP BAY COOL FAULT ALTN FAN INOP
Backup FAN inoperative.
21 EQUIP BAY COOL FAULT SUPP FAN INOP
Supplementary bay fan inoperative.
21 EQUIP BAY COOL FAULT IFAN INOP
FWD or MID Inlet FAN inoperative.
21 EQUIP BAY One channel of either FWD or MID or SUPP bay ventilated sensor failed. COOL FAULT AVIO TEMP SNSR REDUND LOSS 21 EQUIP BAY COOL FAULT FWD AVIO EXHAUST VLV INOP
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FWD AEV failed in position or unknown position.
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Page Intentionally Left Blank
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Table 29: INFO Messages MESSAGE
LOGIC
21 EQUIP BAY COOL FAULT MID AVIO EXHAUST VLV INOP
MID AEV failed in position or unknown position.
21 EQUIP BAY COOL FAULT MID GND VLV INOP
MID equipment bay ground valve failed in position.
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WING-TO-BODY FAIRING VENTILATION SYSTEM GENERAL DESCRIPTION The wing-to-body fairing (WTBF) ventilation system provides ventilation and heat extraction in the forward section of the body fairing to limit the temperature of the WTBF composite structure. The WTBF ventilation system induces secondary airflow to extract heat radiated by the packs and the high-pressure bleed air ducting. Each air conditioning unit (ACU) ram air exhaust has a low-pressure ejector and mixer duct to induce ventilation. The ejector and mixer duct are located just downstream the ACU plenum exhaust and ram air regulating valve (RARV). When the packs are operating, the heat exchanger cooling air is discharged through the RARV into the ram air exhaust. This airflow draws the WTBF compartment air into the ram air exhaust to be discharged overboard.
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CS1_CS3_2150_003
Ram Air Regulating Valve
Figure 88: Wing-to-Body Fairing Ventilation System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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HIGH-PRESSURE PNEUMATIC DUCT AREA COOLING Ram air is used to cool the area around the high-pressure pneumatic duct. Air is tapped off the ram air inlet ducts at the heat exchanger inlet. Two supply tubes feed a piccolo tube that direct air towards the bulkhead area to ensure that components installed in that area are not overheated.
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Piccolo Tube
High-Pressure Pneumatic Duct
Ram Air Duct
Air Cycle Machine Heat Exchanger Inlet
CS1 CS3 2150 031 CS1_CS3_2150_031
Supply Tubes
Figure 89: High-Pressure Pneumatic Duct Area Cooling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21-30 CABIN PRESSURE CONTROL SYSTEM GENERAL DESCRIPTION
MANUAL MODE
The cabin pressure control system (CPCS) automatically controls the pressure inside the cabin to a safe and comfortable level by modulating the flow of pressurized air through the outflow valve (OFV).
The manual mode is selected using the AUTO PRESS PBA. IASC channel B controls the outflow valve through the safety channel, and the MAN RATE rotary switch adjusts the cabin rate of change. In this mode, both automatic channels are inactive. Only one manual channel is in control at a time.
The system consists of two integrated air system controllers (IASCs). Each IASC has three channels: channel A, channel B, and a safety channel. The pressurization system has three independent modes of operation each driving its own actuator on the outflow valve through the CABIN PRESSURE relay. There are two automatic modes controlled by IASC channel A and one manual mode operated from the PRESSURIZATION panel through the IASC safety channel. The manual mode also provides overpressure protection. Auto channel 1 is powered by DC ESS BUS 1, and auto channel 2 is powered by DC ESS BUS 2. The manual channel is powered by the IASC in control. At takeoff, the outflow valve starts to close to pre-pressurize the cabin. During flight, the OFV is normally slightly open in order to maintain the cabin pressure. After landing, the OFV is driven full open to depressurize the aircraft. On the ground the OFV is normally fully open to equalize the inside pressure to the outside ambient pressure to ensure the doors can be opened safely. AUTOMATIC MODE In normal mode, the system operates automatically during all flight phases based on predetermined schedules using IASC channel A. The cabin altitude reference is based on inputs from aircraft systems through the data concentrator unit module cabinets (DMCs). Both IASCs are active in automatic mode, but only one is in control. IASC 1 is in control on odd days, and IASC 2 is in control on even days. If one IASC fails in auto mode, the second IASC automatically takes over. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
DITCHING The DITCHING function is always available, whether the system is in AUTO mode or in MANUAL mode. When the DITCHING PBA is selected, the OFV is driven open to depressurize the aircraft. The packs are shut OFF automatically, and all valves below the flotation line are closed. Prior to ditching, the OFV is driven close to avoid water intake. EMERGENCY DEPRESSURIZATION Emergency depressurization is always available in AUTO or MANUAL mode. When the EMER DEPRESS PBA is selected, the OFV is driven open. The rapid depressurization function has protection against excessive cabin rate of change, and cabin altitude. SAFETY FUNCTIONS Two safety valves (SFVs) protect the pressurized fuselage against overpressure conditions. The SFV open if the cabin differential pressure begins to exceed the structural limits. The SFVs also provide negative pressure-relief. A negative pressure-relief valve (NPRV) provides additional protection against negative differential pressure. Pressure equalization valves (PEV) maintain pressure equilibrium between the pressurized cabin, and the cargo compartment. The large PEV allows the air from cabin into the cargo compartment, while the small PEV allows the airflow out of the cargo compartment into the cabin.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
DITCHING AUTO PRESS FAIL
DMC 1 and DMC 2 IASC 1
ON
MAN
MAN RATE
PAX OXY
CDC 1 DPLY DN
Auto 1
DC ESS BUS 1
UP
PRESSURIZATION PANEL
CABIN PRESSURE RELAY
CDC 2 DC ESS BUS 2
Manual
Auto 2 OUTFLOW VALVE
IASC 2 LEGEND Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Negative Pressure-Relief Valve Pressure Equalization Valve – Large Pressure Equalization Valve – Small Safety Valve AFDX ARINC 429 RS 422
PEV L FWD CARGO COMPARTMENT
SFV 2 PEV S PEV L
AFT CARGO COMPARTMENT
PEV S
NPRV AFT Pressure Bulkhead
CS1_CS3_2130_017
CDC DMC IASC NPRV PEV L PEV S SFV
SFV 1
Figure 90: Cabin Pressure Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
TYPICAL FLIGHT PROFILE
Climb
In normal automatic cabin pressure control mode, a typical flight profile includes the following sequences:
The climb sequence starts 6000 ft above takeoff altitude or 10 minutes after takeoff. The cabin rate of climb is controlled for passenger comfort. The cabin climb profile varies directly with the aircraft climb rate.
•
Prepressurization
•
Takeoff
Cruise
•
Return to base
•
Climb
•
Cruise
The cruise sequence starts when the aircraft rate of climb is less than 500 fpm for 10 seconds. The cabin altitude is maintained in accordance with the pressurization schedule. The cabin altitude is controlled to a maximum of 7840 ft at 41000 ft aircraft altitude.
•
Descent
•
Depressurization
Descent
Prepressurization Prepressurization starts when the aircraft is on the ground and the thrust levers are advanced to the takeoff range. The aircraft pressurizes to 300 ft below the takeoff field elevation to minimize the pressure bump at takeoff. Takeoff When the aircraft leaves the ground, the takeoff sequence is initiated. The takeoff sequence is maintained until the aircraft is more than 6,000 ft above the takeoff field elevation or 10 minutes has elapsed. The main function of the takeoff sequence is to avoid having to reselect the landing altitude in case of a need to return and land at the takeoff field.
The descent sequence starts when the aircraft descends at more than 500 fpm for 10 seconds. Cabin altitude descends following the pressurization schedule until the cabin altitude is 300 ft below the landing field elevation. This minimizes the pressure bump at landing. Depressurization Mode The depressurization sequence starts when the aircraft lands. The cabin pressurizes to the field elevation at a rate of 300 fpm for 45 seconds, and then is driven fully open to minimize any pressure differential that might prevent the doors from opening safely.
Return-to-Base The return-to-base sequence starts if the aircraft descends at more than 500 fpm for more than 10 seconds during the takeoff sequence. The aircraft depressurizes to 300 ft below the landing field elevation.
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21 - Environmental Control 21-30 Cabin Pressure Control System
rc
ra
ft
Al
tit
ud e
41,000 ft
Ca
bi
n
Al
tit
ud
e
Ai
7840 ft
500 ft/min
Takeoff Elevation
300 ft/min
Landing Elevation -300 ft
CS1_CS3_2130_011
-300 ft
Figure 91: Typical Flight Profile Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
COMPONENT LOCATION
PRESSURE EQUALIZATION VALVES
The following components are part of the cabin pressure control system:
The pressure equalization valves (PEVs) are located in the cargo compartment entrance way. The forward cargo compartment PEVs are aft of the cargo door. The aft cargo compartment PEVs are forward of the cargo door.
•
Outflow valve (OFV)
•
Water activated check valve
•
Muffler
•
Safety valve (SFV)
•
Negative pressure-relief valve (NPRV)
•
Pressure equalization valves (PEVs)
OUTFLOW VALVE The outflow valve is located in the environmental control system bay on the bulkhead. WATER ACTIVATED CHECK VALVE The water activated check valve is located in the wing-to-body-fairing (WTBF). MUFFLER The muffler is located in the WTBF. SAFETY VALVE There are two safety valves (SFVs) located on the aft pressure bulkhead. NEGATIVE PRESSURE-RELIEF VALVE The negative pressure-relief valve (NPRV) is located on the aft pressure bulkhead.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Safety Valve Pressure Equalization Valve – Large Pressure Equalization Valve – Small Water Activated Check Valve
Negative Pressure-Relief Valve
CS1_CS3_2130_013
Pressure Equalization Valve – Large Pressure Equalization Valve – Small
Muffler Outflow Valve
Figure 92: Pressurization Components Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
COMPONENT INFORMATION OUTFLOW VALVE The outflow valve modulates the discharge airflow to control the cabin pressure in both AUTO modes and in MANUAL mode. The OFV actuator assembly consists of two identical auto channels and a manual channel. A travel limiter physically limits the outflow valve opening when deployed. When the aircraft altitude exceeds 17,000 ft, the travel limiter pin extends to limit the outflow valve travel.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Outflow Valve Actuator
CS1_CS3_2130_014
Travel Limiter Pin Travel Limiter
Figure 93: Outflow Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
WATER ACTIVATED CHECK VALVE The water activated check valve is a mechanical flapper installed on the environmental control system bay bulkhead. The flapper is supported by a bracket assembly under normal conditions. If the aircraft ditches, the rising water level closes the flapper. MUFFLER The optional muffler reduces the outflow valve discharge air noise level.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Bracket Assembly
Water Activated Check Valve Pressure Bulkhead
Muffler
CS1_CS3_2130_007 S1 CS3 2130 007
Water Activated Check Valve
Figure 94: Outflow Valve, Water Activated Check Valve, and Muffler Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
SAFETY VALVE Two safety valves protect against any overpressure condition. The safety valves are pneumatically-actuated on the ground or in flight. The SFVs provide protection for both overpressure relief, and negative pressure-relief. NEGATIVE PRESSURE-RELIEF VALVE The negative pressure-relief valve (NPRV) is a mechanical check valve with spring-loaded flappers. The NPRV is acted on by ambient pressure on one side, and cabin pressure on the other side to sense the differential pressure.
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21 - Environmental Control 21-30 Cabin Pressure Control System
A
B
B NEGATIVE PRESSURE-RELIEF VALVE
A SAFETY VALVE AFT PRESSURE BULKHEAD
CS1_CS3_2130_009
Flappers
Figure 95: Safety Valves and Negative Pressure-Relief Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURE EQUALIZATION VALVES The system maintains pressure equalization between the cabin and forward and aft cargo using four pressure equalization valves (PEVs). The PEV is a fully mechanical device that prevents excessive pressure between the cargo compartment and cabin underfloor in both directions (inflow and outflow). Two PEVs are installed on each cargo compartment entryway. Each cargo compartment has one large cargo PEV, and one small PEV. The large PEV allows the air from the cabin into the cargo compartment, and the small PEV allows the airflow out of the cargo compartment into the cabin. The large PEV prevents the cabin from pressurizing when the cargo doors are not safely closed, and prevents the cargo blowout panels from opening. The small PEV allows cargo compartment pressure to equalize pressure with cabin pressure during fast climb.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Pressure Equalization Valve – Large
NOTE Aft cargo shown. Fwd cargo similar.
CS1_CS3_2130_008
Pressure Equalization Valve – Small
Figure 96: Pressure Equalization Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
DETAILED COMPONENT INFORMATION INTEGRATED AIR SYSTEM CONTROLLER The system consists of two identical integrated air system controllers (IASCs). Each IASC incorporates channel A and channel B and a safety channel. A cabin pressure sensor is fitted to each channel A and channel B of the IASC so that the cabin pressure control system receives four independent cabin pressure signals. The safety channel receives cabin pressure information from the channel B sensor. The safety channel limits cabin altitude and cabin rate of change, and sends cabin altitude and cabin pressure rate of change to the data concentrator unit module cabinet (DMC) by an ARINC 429 BUS.
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21 - Environmental Control 21-30 Cabin Pressure Control System
IASC 1
Channel A
Channel B
Safety Channel
• Automatic pressure control
• Manual pressure control
• Cabin altitude limitation
• Emergency depressurization mode
• Emergency depress mode
• Ram air (OFV forced open)
• Ditching sequence
• Ditching sequence
• Manual pressure control
PS
• Cabin altitude limitation
PS
• Cabin pressure sensor
Data Exchange
DMC
IASC 2
Channel B
Safety Channel
• Automatic pressure control
• Manual pressure control
• Cabin altitude limitation
• Emergency depressurization mode
• Emergency depress mode
• Ram air (OFV forced open)
• Ditching sequence
• Ditching sequence
• Manual pressure control
PS
• Cabin altitude limitation
PS
ARINC 429 DMC Data Concentrator Unit Module Cabinet OFV Outflow Valve PS
Pressure Sensor
• Cabin pressure sensor CS1_CS3_2190_005
LEGEND
Channel A
Figure 97: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
CONTROLS AND INDICATIONS PRESSURIZATION PANEL The PRESSURIZATION panel has controls for: •
Automatic cabin pressure control
•
Manual cabin pressure control
•
Rapid cabin pressure-relief (EMER DEPRESS)
•
Ditching
The system is normally in AUTO mode for cabin pressurization. If the cabin pressure AUTO fails, the amber FAIL light illuminates. Selecting the AUTO PRESS PBA to MAN, switches the cabin pressure control mode to MANUAL. In manual mode, the MAN RATE knob adjusts the cabin altitude. Counterclockwise selection decreases the cabin altitude at a maximum rate of -2,500 ft/min. Clockwise selection increases the cabin altitude at a maximum rate of +2,500 ft/min. There are detents at the +1,000 ft positions on the switch. Selecting the guarded EMER DEPRESS PBA to ON, opens the outflow valve. The cabin altitude and rate limitations functions (15,000 ft) remain active. The white OPEN legend indicates the outflow valve is being forced open. The DITCHING PBA closes all of the valves mounted on the aircraft located below the flotation line, in preparation for a ditching. Selecting the guarded DITCHING PBA opens the outflow valve when altitude is below 15,000 ft. All valves are closed electrically. Prior to ditching, the outflow valve is driven closed.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
DITCHING AUTO PRESS FAIL
ON
MAN
MAN RATE
PAX OXY
DPLY UP
PRESSURIZATION PANEL
CS1_CS3_2130_004
DN
Figure 98: Pressurization Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
FMS PAGE The landing elevation is selected via the flight management system (FMS) through an automatic mode or two backup manual modes. When the flight plan is programmed in the FMS ROUTE tab, the landing elevation of arrival airport data is displayed and also sent via an ARINC 429 BUS to the IASCs. The landing elevation is also set using: •
Control tuning panel (CTP)
•
AVIONIC CTP page
Using the CTP, the LDG ELEV left line select key switches between FMS and MAN, and the LDG ELEV right line select key changes the MAN elevation. The CTP input can be overridden using the avionics virtual CTP FMS/MAN LDG ELEV selection to set the landing elevation manually. In either manual mode, the landing elevation is sent to the IASCs via an ARINC 429 BUS.
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21 - Environmental Control 21-30 Cabin Pressure Control System
BRT
HDG
MAG/TRUE
OFF
TUNE/ MENU
11234
AIR
DOOR
ELEC
FLT CTR CTRL TRL L
FUEL FUE EL
HYD YD
AVIONIC
INFO
CB
AVIOO
IDENT
BRG 1/2>
LDG ELEV FMS/MAN
STATU STATUS TATUS S
L CTP OVRD 850 FT
TUNE/DATA
NAV SRC
FMS 1 BARO
ACT FPLN 24R
NAV SRC
000
FMS 1 BARO
UNITS
FINAL APPR CRS
108.50
251T
RWY THRESHOLD ELEV 126 FT
GS ANGLE
3.00°
MINIMUMS
UNITS
MINIMUMS
OFF BRG 2
BRG 1
BRG 2
OFF
OFF
OFF
OFF
HDG
FPV CAGE
HDG
FPV CAGE
MAG TRUE
ON OFF
MAG TRUE
ON OFF
FMS MAN
YES
000
BRG 1
LDG ELEV
WGS-84
CRS
00.00 STD IN
OFF
FREQUENCY
ARRIVALS...
CRS
00.00 STD IN
CONTROL TUNING PANEL (CTP)
Landing Elevation Normally Set on Arrival Page in FMS
R CTP OVRD
1/2
RETURN
ARRIVAL DATA KLAX ILS
CTP TP
850
FT
LDG ELEV
FMS MAN
DONE DON NE SYNOPTIC PAGE – AVIONICS
FMS ARRIVAL PAGE
000
FT CS1_CS3_2130_005
MINIMUMS BARO/RAD
Figure 99: FMS, Control Tuning Panel, and AVIONIC CTP Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
INDICATIONS The cabin pressure control system information is displayed on the engine indication and crew alerting system (EICAS) page or the AIR synoptic page. The CAB ALT value is the highest recorded value by any of the six IASC channels. If the landing elevation is set manually, MAN appears next to the LDG ELEV value. An arrow is displayed beside the RATE to indicate the direction of the cabin rate of climb. •
If the cabin rate of climb increases, an upward white arrow appears
•
If the cabin rate of climb decreases, a downward white arrow appears.
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21 - Environmental Control 21-30 Cabin Pressure Control System
TOTAL FUEL (KG) 10905 2735
1
5435
5500 8.0 ELEV 560
2735
CAB ALT
RATE
ó3
CREW
LDG 7(03 (°C)
23
100 OXY 2000
TRIM NU STAB
LO
22
22
CAB ALT Description
Value
4.2
ND
NL RUDDER NR
INFO
XXXXX
Cabin altitude above warning threshold
XXXX or XXXXX
Cabin altitude above caution threshold
XXXXX
Cabin altitude within normal range
------
Invalid value
Value
ǻ3 Description
Value
36, 36,
XXXXX
Cabin altitude manually selected or pressurization system in manual mode
XXXXX
Cabin altitude actual value
-----
Invalid value
3
(,&$63$*(
5 -45
RATE Description
X.X
Delta P above warning threshold
XX.X
Delta P within normal range
--.-
Invalid value
Value
560
5 -45
XBLEED
36,
560 MAN -----
LDG ELEV Description LDG ELEV selected from FMS LDG ELEV selected through CTP or AVIONIC synoptic page Invalid value
$38
ó3 LDG ELEV
5500 8.0 560
6<1237,&3$*(±$,5
RATE CREW OXY
100 2000
1
Air block not shown in EICAS compressed mode.
2
Air block only shown on SYNOPTIC page when EICAS in compressed mode.
3
The cabin rate arrow and digital readout in cyan for 3 seconds when manually turns selected.
2
CS1_CS3_2130_006
NOTE
CAB ALT
Figure 100: EICAS and AIR Synoptic Page Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
DETAILED DESCRIPTION AUTO/MANUAL SELECTION The system is normally in AUTO mode for cabin pressurization. If the IASC controlling the automatic pressurization fails, the standby channel automatically takes over. If both IASC automatic channels fail, the AUTO PRESS PBA indicates FAIL, and an AUTO PRESS FAIL caution message is displayed. When the AUTO PRESS PBA is selected to MAN, the cabin pressure control mode switches to manual. The AUTO PRESS PBA MAN indication illuminates white and a CABIN PRESS MAN status message is displayed. The AUTO PRESS PBA can also be used to select active channels if required. When in automatic mode, pressing the AUTO PRESS PBA selects the manual channel of the IASC. A second press of the pushbutton annunciator (PBA) selects the other IASC automatic channel. A third press of the PBA selects the other IASC manual channel.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION
Auto 1 IASC 1
EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
Manual IASC 1
AUTO PRESS FAIL MAN
Auto 2 IASC 2
MAN RATE
DPLY DN
UP P
Manual IASC 2
CS1_CS3_2130_003
PRESSURIZATION PANEL
Figure 101: Cabin Pressure Control System Selection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
INTERFACE The outflow valve (OFV) actuator assembly consists of two identical auto channels and a manual channel. The auto channels of the OFV communicate with their corresponding integrated air system controller (IASC) channel A via an RS 422 BUS. The manual channel is hardwired to the IASCs through the CABIN PRESSURE relay, located in electrical power center 2 (EPC 2). The CABIN PRESSURE relay allows both IASCs to manually command the OFV. Only one controller at a time is connected to the MANUAL channel of the OFV. The outflow valve commands are provided by the IASC in control. The outflow valve manual channel provides position feedback to each IASC.
The aircraft cannot be dispatched if the travel limiter pin is failed closed. If the pin fails in the extended position, the aircraft can be dispatched but is limited to a maximum altitude of 25,000 ft. Cabin Altitude Limitation The cabin altitude limitation system limits cabin altitude to a maximum of 15 000 ft. When the altitude limit is reached, a signal is sent through the CABIN PRESSURE relay to CDC 1 and CDC 2. The CDCs remove power from the auto channels of the OFV. The outflow valve is driven closed using the manual channel of the IASC that was in control.ever
When the aircraft is on ground and the doors are open, the IASCs perform a MANUAL channel test as follows: •
The valve is moved from the fully open position to the fully closed position in 5 seconds
•
The valve is moved back to the fully open position in 5 seconds
•
If the valve position feedback does not match the commanded valve position, the MANUAL channel is reported as failed
A travel limiter device physically limits the outflow valve opening when extended. The travel limiter extends based on a preset cabin differential pressure, either in automatic or manual modes, when the aircraft flies over 17,000 ft. The travel limiter is powered by IASC 1. The travel limiter provides position feedback to IASC 1. The active IASC channel A software continuously monitors for any overpressure condition. If an overpressure condition is detected, the active IASC drives the OFV open. A force open command signal is generated for an emergency depressurization, or when the emergency ram air valve is opened. The force open command removes power from the auto channels, and drives the OFV open using the manual channel.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Cabin Pressure Relay Active
OFV Force Open Command
LEGEND CABIN PRESSURE RELAY
CDC EPC IASC OFV
Control and Distribution Cabinet Electrical Power Center Integrated Air System Controller Outflow Valve RS 422
Altitude Limit Auto 2 Manual Manual Close Manual Close Manual Open Manual Open Manual Position IASC 2
EPC 2 Cabin Pressure Relay Active Altitude Limit Manual Close Manual Open
CDC 2-6-12 Outflow VLV MAN DC ESS SSPC 3A BUS 2 Logic: Combinational Altitude Limit CDC 2-5-9 Outflow VLV AUTO 2 DC ESS SSPC 3A BUS 2
Manual Position
Logic: Combinational
CDC 2 Travel Limiter
Manual Position
Travel Limiter
Travel Limiter Position
Travel Limiter Position
FOV Force Open Command IASC 1
CDC 1-6-9 Outflow VLV AUTO 1 DC ESS SSPC 3A BUS 1 Logic: Combinational Altitude Limit
Auto 1 OFV
CDC 1
CS1_CS3_2130_015
Travel Limiter
Figure 102: Cabin Pressure Control System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
CABIN ALTITUDE LIMITATION ANNUNCIATION
CABIN ALTITUDE WARNING LOGIC
The baro-corrected landing elevation (L/EBARO) has three modes of operation for the cabin altitude limitation. One is for normal operation, and the other modes are for high altitude airport operations (HAAO).They are defined as follows:
The normal CAB ALT warning message threshold is 9865 ft.
•
Non-HAAO conditions defined for L/EBARO between -2000 ft and 8000 ft above sea level.
•
HAAO-1 conditions defined for L/EBARO between 8000 ft and 14,000 ft above sea level
•
HAAO-2 conditions are defined for L/EBARO between 14,000 ft and 14,500 ft above sea level
The cabin altitude caution and warning EICAS messages for non-HAAO operations, are set at: •
CAB ALT CAUTION 8500 ft
•
CAB ALT WARNING: 9865 ft
If the landing altitude is above 8000 ft: •
When aircraft starts the descent, the CABIN ALTITUDE warning starts to increase as a function of cab altitude caution threshold + 1365 ft and limited to 14,500 ft in HAAO-1 or 14,843 ft in HAAO-2.
If the takeoff is above 8,000 ft: •
When aircraft starts its climb and if aircraft altitude is above TOA + 5000 ft or 10 minutes after takeoff, the cabin altitude warning starts to decrease as a function of aircraft altitude to it nominal cruise set point of 9865 ft
In either landing or take off at high altitude, a CABIN ALT LEVEL HIGH advisory message is displayed to indicate that the cabin altitude warning threshold has been set above 9865 ft. The system limits cabin altitude to a maximum of 14843 ft in non-HAAO or HAAO-1 and 15200 ft in HAAO-2 in the event of a system malfunction.
CABIN ALTITUDE CAUTION LOGIC The normal CAB ALT CAUTION threshold is 8500 ft. If the landing altitude is above 8000 ft: •
When aircraft starts the descent and the aircraft altitude is below 37,000 ft, the CABIN ALTITUDE caution starts to increase as a function of aircraft altitude to a maximum of landing altitude + 500 ft and limited to 14,500 ft in HAAO-1 OR 14,843 ft in HAAO-2
The caution threshold increase rate is proportional to the A/C descent rate, aircraft altitude and L/EBARO. If the takeoff is above 8000 ft: •
When aircraft starts its climb and if aircraft altitude is above takeoff altitude (TOA) + 5000 ft or 10 minutes after takeoff, the CABIN ALTITUDE caution starts to decrease as a function of aircraft altitude to it nominal cruise set point of 8500 ft
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21 - Environmental Control 21-30 Cabin Pressure Control System
Caution Altitude
TOA Takeoff Altitude
CRUISE 41,000 ft
CRUISE 41,000 ft
Start of Cabin Climb
CAB ALT Warning TOA +1865 ft CAB ALT Caution TOA +500 ft 300 ft CAB ALT Warning 9865 ft
HAAO Landing Elevation
CAB ALT Warning TOA +1865 ft TOA CAB ALT Caution TOA +500 ft Takeoff Altitude 300 ft > 8000 ft
5000 ft
10 minutes
CAB ALT Caution 8500 ft
CAB ALT Warning 9865 ft CAB ALT Caution 8500 ft
Cabin Altitude Max 8000 ft
Cabin Altitude Max 8000 ft
Cruise
HAAO Takeoff Sequence
HAAO Descent
Climb
Cruise
Time
Time HAAO TAKEOFF CAB ALT LOGIC CS1_CS3_2130_012
HAAO LANDING CAB ALT LOGIC
Figure 103: Cabin Altitude Limitations Annunciation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
EMERGENCY DEPRESSURIZATION The cabin pressure can be reduced rapidly towards ambient pressure in an emergency. This function is known as emergency depressurization. Emergency depressurization is always available, whether the system is in AUTO mode, or in MANUAL mode. While in AUTO mode at the time of the emergency depressurization selection, a cabin rate of change limitation of 3500 ft/min is applied. In the MANUAL mode, the cabin rate of change limitation of 2500 ft/min is applied. The cabin altitude is limited to 14,500 ft. When the EMER DEPRESS PBA is pressed, the OFV is driven open at a rate that protects against excessive cabin rate of change. The ON legend on the PBA indicates the OFV is being driven open. An EMER DEPRESS ON caution message is displayed on EICAS. If the system was in automatic mode at the time of emergency depressurization selection, the function is accomplished by the active IASC channel A. If the system was in manual mode at the time of the emergency depressurization selection, a force open command is supplied through control and distribution cabinet 1 (CDC 1) to the safety channel. The outflow valve is driven open through the manual channel. The cabin altitude and rate limitations functions remain active. The CAB ALT caution and warning are also raised above the normal set point in order to monitor that the cabin altitude does not exceed 15,000 ft.
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21 - Environmental Control 21-30 Cabin Pressure Control System
PRESSURIZATION EMER DEPRESS MAX ∆P 0.11 PSI TO 1.00 PSI LDG
ON
DMC 1 and DMC 2
DITCHING AUTO PRESS FAIL
Channel A
Channel A
ON
MAN
MAN RATE
PAX OXY
DPLY DN
Channel B
UP
CDC 1
Channel B
CDC 2
PRESSURIZATION PANEL
Force Open Command Safety Channel
Force Open Command Safety Channel
CABIN PRESSURE RELAY
IASC 1
IASC 2
CDC DMC IASC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller AFDX ARINC 429 RS 422
Auto 2
Manual OUTFLOW VALVE
Auto 1
CS1_CS3_2130_018
LEGEND
Figure 104: Emergency Depressurization Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
DITCHING The ditching mode is used to close all of the valves below the aircraft flotation line. Ditching can be selected in either AUTO or MANUAL pressurization modes. The DITCHING PBA closes all valves in preparation for a water landing. When the aircraft is below 15,000ft, pressing the guarded DITCHING PBA:
The following valves prevent water intake mechanically
•
Opens the outflow valve
•
Closes the flow control valves
The following valves are above the flotation line, but prevent water intake:
•
Trim air pressure-regulating valve
•
Safety valves
•
Trim air shutoff valve
•
Negative pressure-relief valve
•
Low-pressure ground connection
•
Bulkhead check valves
•
Water activated check valve
Closes all electrically controlled valves located below the flotation line •
Avionics exhaust valves
•
Ground valve
When pressing the DITCHING PBA, it latches and the ON legend illuminates white to indicate the electrically controlled valves have closed. If the ram air valve is open, the RAM AIR PBA is pressed to close it. A DITCHING MISCONFIG caution message indicates the DITCHING PBA is selected ON, and the ram air valve is open. When the cabin differential pressure has equalized, all valves are closed, and the airspeed is less than 80 kt, the outflow valve is driven closed. A DITCHING ON status message indicates the DITCHING PBA is pressed and the ditching sequence is complete.
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21 - Environmental Control 21-30 Cabin Pressure Control System
SFV (2X) – NPRV
AEV
GV
TAPRV TASOV
ERAV FCV
WACV LPGC
OFV
BCKV
AEV
ELECTRICALLY CLOSED VALVES
MECHANICALLY CLOSED VALVES
Forward AEV
BCKV
OFV
WACV
GV
LPGC
Mid AEV
SFV
FCV, TAPRV, TASOV
NPRV
AEV BCKV ERAV FCV GV LPGC NPRV SFV TAPRV TASOV WACV
Avionics Exhaust Valve Bulkhead Check Valve Emergency Ram Air Valve Flow Control Valve Ground Valve Low-Pressure Ground Connection Negative Pressure-Relief Valve Safety Valve Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Water Activated Check Valve
ERAV
CS1_CS3_2130_016
LEGEND
Figure 105: Ditching Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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21 - Environmental Control 21-30 Cabin Pressure Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the cabin pressure control system:
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21 - Environmental Control 21-30 Cabin Pressure Control System
CAS MESSAGES
Table 32: ADVISORY Messages Table 30: WARNING Messages
MESSAGE
MESSAGE
LOGIC
CABIN DIFF PRESS
Positive differential pressure is above pneumatic relief setting.
CABIN ALT
Cabin altitude exceeds 10,000 ft under non-HAAO conditions, and up to 15,000 ft under HAAO conditions posted by IASC A channels. Cabin altitude exceeds 15,000 ft under any condition posted by IASC B channels. Safety cabin altitude high alarm from IASC1C or IASC2C.
Table 31: CAUTION Messages MESSAGE CABIN ALT
LOGIC No CABIN ALTITUDE warning message and cabin altitude exceeds 8,500 ft under non-HAAO conditions, and up to 15,000 ft under HAAO conditions. Cabin altitude exceeds 10,000 ft under HAAO conditions for more than 30 minutes.
AUTO PRESS FAIL
Both cabin pressure control AUTO functions are failed.
DITCHING MISCONFIG
DITCHING selected ON and RAM AIR selected OPEN.
LDG ELEV MISCONFIG
High airport altitude landing is selected and HAAO option is not present on aircraft.
EMER DEPRESS ON
EMER DEPRESS selected ON confirmed by either one auto mode or one manual mode.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
CABIN ALT LEVEL High airport altitude landing or takeoff sequence is HI initiated and the system operates in automatic mode. (increase of caution and warning threshold above normal setting). PRESSURIZATION FAULT
Loss of redundant or non-critical function for the pressurization system.
Table 33: STATUS Messages MESSAGE
LOGIC
DITCHING ON
DITCHING selected ON and sequence completed.
CABIN PRESS MAN
MANUAL PRESSURIZATION mode selected.
Table 34: INFO Messages MESSAGE
LOGIC
21 PRESSURIZATION FAULT - PRIM ALT LIM INOP
Loss of IASC 1 altitude limitation function.
21 PRESSURIZATION FAULT - BACKUP ALT LIM INOP
Loss of IASC 2 altitude limitation function.
21 PRESSURIZATION FAULT - CPCS AUTO MODE REDUND LOSS
One automatic mode failed.
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21 - Environmental Control 21-30 Cabin Pressure Control System
Table 34: INFO Messages MESSAGE
LOGIC
21 PRESSURIZATION FAULT - MANUAL MODE INOP
OFV manual mode failed.
21 PRESSURIZATION FAULT - OFV FINGER FAIL OUT
Altitude limiter failed deployed.
21 PRESSURIZATION FAULT - OFV FINGER FAIL IN
Altitude limiter failed in closed position.
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ATA 28 - Fuel
BD-500-1A10 BD-500-1A11
28 - Fuel
Table of Contents 28-00 Fuel Safety ...................................................................28-2 Fuel safety History ............................................................28-2 25.981 Fuel Tank Ignition Prevention ..........................28-2 Special Federal Aviation Regulation 88 .......................28-2 European Aviation Safety Agency Regulations ...........28-2 Airworthiness Limitations-Fuel System Limitations ..........28-4 Critical Design Configuration Control Limitations .............28-6 CDCCL Warning ...............................................................28-8 Fuel Tank Access and Venting .......................................28-10 Fuel Tank Access ......................................................28-10 Fuel Tank Venting......................................................28-10 Fuel Tank Entry ..............................................................28-12 Preparation ................................................................28-12 Entry...........................................................................28-12 28-11 Fuel Storage ...............................................................28-14 General Description ........................................................28-14 Component Location .....................................................28-16 Flapper Check Valves................................................28-16 Fuel Drain Valves.......................................................28-16 Fuel Tank Access Panels ..........................................28-16 Practical Aspects ............................................................28-18 Fuel Drain Valve Servicing.........................................28-18 28-12 Vent System ...............................................................28-20 General Description .......................................................28-20 Component Location ......................................................28-22 Float Drain Valves......................................................28-22 Center Tank Float Vent Valves ..................................28-22 Wing Tank Float Vent Valves.....................................28-24 Bifurcated Vent Openings ..........................................28-24 NACA Inlet Scoop ......................................................28-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Vent Relief Flame Arrestor ........................................ 28-26 Positive Pressure-Relief Valves ................................ 28-28 Center Tank Positive Pressure-Relief Valve ............. 28-28 Center Tank Negative Pressure-Relief Valve ........... 28-28 28-23 Refuel/Defuel System ................................................ 28-30 General Description ....................................................... 28-30 Refueling ................................................................... 28-32 Defueling ................................................................... 28-36 Component Location .................................................... 28-40 Refuel Adapter .......................................................... 28-40 Refuel/Defuel Control Panel...................................... 28-40 Refuel Shutoff Valve ................................................. 28-42 Refuel Control Solenoid Valve .................................. 28-42 Defuel/Isolation Transfer Shutoff Valve and Actuator ....................................... 28-44 Manual Defuel Shutoff Valve and Lever.................... 28-44 Component Information ................................................. 28-46 Refuel Adapter .......................................................... 28-46 Refuel/Defuel Control Panel...................................... 28-48 Controls and Indications ............................................... 28-50 Refuel/Defuel Control Panel...................................... 28-50 Optional Virtual Refuel Panel .................................... 28-54 Operation ....................................................................... 28-56 Auto Refuel Operation............................................... 28-56 Manual Refuel Operation .......................................... 28-58 Pressure Defuel Operation........................................ 28-60 Suction Defuel Operation .......................................... 28-62 Detailed Description ..................................................... 28-64 Refuel Power............................................................. 28-64 Practical Aspects ........................................................... 28-66 Onboard Maintenance System.................................. 28-66
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28 - Fuel
28-21 Engine Feed System ..................................................28-70 General Description .......................................................28-70 Component Location .....................................................28-72 Engine Feed Ejector Pump ........................................28-72 AC Boost Pump .........................................................28-72 Engine Feed Shutoff Valve and Actuator...................28-72 Engine Feed Pressure Switch....................................28-72 Component Information .................................................28-74 AC Boost Pump .........................................................28-74 Controls and Indications ................................................28-76 Engine Fire PBA ........................................................28-76 Fuel Panel..................................................................28-76 Synoptic Display ........................................................28-78 Detailed Description ......................................................28-80 Left Engine Feed........................................................28-80 Right Engine Feed .....................................................28-82 Monitoring and Tests .....................................................28-84 CAS Messages ..........................................................28-85 28-21 APU Feed System ......................................................28-86 General Description .......................................................28-86 Component Location ......................................................28-88 APU Feed Shutoff Valve and Actuator.......................28-88 APU Feed Line, Shroud, and Drain Masts.................28-90 Controls and Indications .................................................28-92 APU Start Switch .......................................................28-92 APU Fire PBA ............................................................28-92 Detailed Description ......................................................28-94 Monitoring and Tests .....................................................28-96 CAS Messages ..........................................................28-97
Automatic Transfer ................................................. 28-102 Center Tank to Main Tank Fuel Transfer ................ 28-104 Manual Fuel Transfer .............................................. 28-106 Component Location ................................................... 28-112 Scavenge Ejector Pump.......................................... 28-112 Transfer Float Valve................................................ 28-114 Transfer Ejector Pump ............................................ 28-114 Gravity Crossflow Shutoff Valve.............................. 28-116 Controls and Indications ............................................. 28-118 Synoptic Page ......................................................... 28-120 Monitoring and Tests .................................................. 28-122 CAS Messages ....................................................... 28-123 28-40 Fuel Indicating System ............................................ 28-124 General Description .................................................... 28-124 Component Location .................................................. 28-126 Fuel Probes............................................................. 28-126 Temperature Sensors ............................................. 28-126 Remote Data Concentrator ..................................... 28-128 Fuel Quantity Computer .......................................... 28-130 Fuel Probes............................................................. 28-132 Temperature Sensor ............................................... 28-132 Detailed Component Information ................................ 28-134 Fuel Quantity Computer Interface ........................... 28-134 Controls and Indications ............................................. 28-136 Detailed Description ................................................... 28-138 Monitoring and Tests .................................................. 28-140 CAS Messages ....................................................... 28-141 Practical Aspects ........................................................ 28-144 Onboard Maintenance System ............................... 28-144
28-22 Fuel Transfer System .................................................28-98 General Description .......................................................28-98 Scavenge System ....................................................28-100 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel
List of Figures Figure 1: Fuel Safety History ................................................28-3 Figure 2: Airworthiness Limitations-Fuel System Limitations ............................................................28-5 Figure 3: Critical Design Configuration Control Limitation ..................................................28-7 Figure 4: Critical Design Configuration Control Limitation Warning in Aircraft Maintenance Publication ..........................28-9 Figure 5: Fuel Tank Access and Venting............................28-11 Figure 6: Fuel Tank Entry ...................................................28-13 Figure 7: Fuel Storage System...........................................28-15 Figure 8: Fuel Storage System Component Location.........28-17 Figure 9: Fuel Drain Valve Servicing ..................................28-19 Figure 10: Fuel Vent System ................................................28-21 Figure 11: Float Drain Valve and Center Tank Float Vent Valve ..................................................28-23 Figure 12: Bifurcated Vent Openings and Wing Tank Float Vent Valve................................28-25 Figure 13: NACA Inlet Scoop and Vent Relief Flame Arrestor.....................................................28-27 Figure 14: Positive Pressure-Relief Valves, Center Tank Positive, and Center Tank Negative Pressure-Relief Valves.........................28-29 Figure 15: Refuel/Defuel System..........................................28-31 Figure 16: Auto Refueling.....................................................28-33 Figure 17: Manual Refueling ................................................28-35 Figure 18: Suction Defueling ................................................28-37 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 19: Pressure Defueling ............................................. 28-39 Figure 20: Refuel Adapter and Refuel/Defuel Control Panel Location................. 28-41 Figure 21: Refuel Control Solenoid Valve and Refuel Shutoff Valve Component Location......... 28-43 Figure 22: Defuel System Components .............................. 28-45 Figure 23: Refuel Adapter.................................................... 28-47 Figure 24: Refuel/Defuel Control Panel ............................... 28-49 Figure 25: Refuel/Defuel Control Panel ............................... 28-51 Figure 26: Refuel/Defuel Control Panel Messages.............. 28-53 Figure 27: Virtual Refuel Panel ............................................ 28-55 Figure 28: Automatic Refuel Operation................................ 28-57 Figure 29: Manual Refuel Operation.................................... 28-59 Figure 30: Pressure Defuel Operation ................................. 28-61 Figure 31: Suction Defuel Operation.................................... 28-63 Figure 32: Refuel Power (Normal Operation - Door Closed)...................... 28-65 Figure 33: Refuel Solenoid Command Test......................... 28-67 Figure 34: Defuel Shutoff Valve Command Test ................. 28-69 Figure 35: Engine Feed System .......................................... 28-71 Figure 36: Engine Feed Component Location ..................... 28-73 Figure 37: AC Boost Pump .................................................. 28-75 Figure 38: Engine Fire PBAs and Engine Fuel Panel .......... 28-77 Figure 39: Fuel Synoptic Page............................................. 28-79 Figure 40: Left Engine Feed Detailed Description ............... 28-81 Figure 41: Right Engine Feed Detailed Description............. 28-83
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Figure 42: APU Fuel Feed System.......................................28-87 Figure 43: APU Feed Shutoff Valve and Actuator ................28-89 Figure 44: APU Fuel Line, Shroud and Drain Masts ............28-91 Figure 45: APU Start Switch and Fire PBA ..........................28-93 Figure 46: APU Feed Shutoff Valve Detailed Description ............................................28-95 Figure 47: Fuel Transfer System ..........................................28-99 Figure 48: Scavenge System .............................................28-101 Figure 49: Automatic Fuel Transfer ....................................28-103 Figure 50: Center Tank to Main Tank Fuel Transfer ..........28-105 Figure 51: Main Tank to Main Tank Fuel Transfer .............28-107 Figure 52: Main Tank to Center Tank Fuel Transfer ..........28-109 Figure 53: Gravity Fuel Transfer.........................................28-111 Figure 54: Scavenge Ejector Pump....................................28-113 Figure 55: Center Tank to Main Tank Fuel Transfer Ejector Pump and Float Valve ............28-115 Figure 56: Gravity Crossflow Shutoff Valve........................28-117 Figure 57: Fuel Transfer Controls.......................................28-119 Figure 58: Fuel Synoptic Page ...........................................28-121 Figure 59: Fuel Quantity Indication.....................................28-125 Figure 60: Fuel Quantity Probes and Temperature Sensors........................................28-127 Figure 61: Remote Data Concentrator ...............................28-129 Figure 62: Fuel Quantity Computer ....................................28-131 Figure 63: Fuel Probes and Temperature Sensor ..............28-133 Figure 64: Fuel Quantity Interface ......................................28-135 Figure 65: Fuel Controls and Indications............................28-137 Figure 66: Fuel Quantity System ........................................28-139 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 67: Low-Level Discrete Output Test ....................... 28-145 Figure 68: Clear NVM Function ......................................... 28-147
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FUEL - CHAPTER BREAKDOWN Fuel Chapter Breakdown
Fuel Safety
Engine Feed System
1
Fuel Storage
5
APU Feed System
2
Vent System
6
Fuel Transfer System
3
Pressure Refuel/Defuel System
4
7
Fuel Indicating System
8
28 - Fuel
28-00 FUEL SAFETY FUEL SAFETY HISTORY Since 1959 there have been 17 fuel tank ignition events, resulting in 542 fatalities, 11 hull losses, and 3 others with substantial damage to the aircraft. Since the 1960's, there have been 5 key accidents involving fuel tank explosions which call into question the fundamental safety strategy applied to fuel systems of large commercial airplanes. Because of this accident, multiple airworthiness directives (ADs) on several aircraft types were issued. 25.981 FUEL TANK IGNITION PREVENTION This advisory circular (AC) 25.981 fuel tank ignition prevention provides guidance for demonstrating compliance with the certification requirements for prevention of ignition sources within the fuel tanks of transport category airplanes. The TWA 800 accident brought the realization that some fuel tanks could be flammable for much of their operational time. In their efforts to determine the cause of TWA 800 accident, the FAA concluded that many current airplanes had problems in their ignition prevention approaches. An additional independent layer of protection would be needed to back up the ignition prevention strategy. SPECIAL FEDERAL AVIATION REGULATION 88
A flammability reduction program would be necessary to prevent the operation of transport category airplanes with explosive fuel-air mixtures in the fuel tank. These fuel tanks were called high flammability tanks. The overall goal was to make sure flammability reduction of high flammability tanks became an integral part of aircraft safety. EUROPEAN AVIATION SAFETY AGENCY REGULATIONS A similar regulation had been recommended by the European Joint Aviation Authorities (JAA) to the European national aviation authorities in JAA letter 04/00/02/07/03-L024 of February 3, 2003. The review was mandated by European national airworthiness authorities using JAR 25.901(c), and 25.1309. In August 2005, the European Aviation Safety Agency (EASA) published a policy statement on the process for developing instructions for maintenance and inspection of fuel tank system ignition source prevention that also included the EASA expectations with regard to compliance times of the corrective actions. The EASA policy statement reset the date for compliance of unsafe items to July 1, 2006. Fuel airworthiness limitations are items from a systems safety analysis that have been shown to have failure modes associated with an unsafe condition as defined in SFAR 88. These are identified in failure conditions for which an unacceptable probability of ignition risk could exist if specific tasks and/or practices are not performed in accordance with the manufacturers requirements.
In response to these findings, the FAA issued Special Federal Aviation Regulation (SFAR) No. 88 in June of 2001. SFAR 88 required a safety review of the aircraft fuel tank system to determine that the design meets the requirements of FAR 25.901 and 25.981 (a) and (b).
This EASA airworthiness directive mandated the fuel airworthiness limitation critical design configuration control limitations (CDCCLs) consisting of maintenance and inspection tasks for the type of aircraft, that resulted from the design reviews, JAA recommendations, and the EASA policy statement.
The safety reviews were very valuable in that they revealed unexpected ignition sources. There was difficulty in identifying all ignition sources and a number of previously unknown failures were found.
The FAA and EASA have also mandated that operators must train their maintenance and engineering personnel regarding the changes brought about by SFAR 88.
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28 - Fuel
Aircraft
Location
Year
Cause
Boeing 707
Elkton, Maryland
1963
Lightning
Boeing 747
Madrid, Spain
1976
Lightning
Boeing 737
Manila, Philippines
1990
Unconfirmed
Boeing 747
New York, New York
1996
Center fuel tank explosion. Exact cause undetermined.
Boeing 737
Bangkok, Thailand
2001
Center fuel tank explosion. Possible center fuel pump failure.
TWA 800 B747
CS1_CS3_2800_010
MAJOR ACCIDENTS RELATED TO FUEL TANK EXPLOSIONS
Figure 1: Fuel Safety History Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-00 Fuel Safety
AIRWORTHINESS LIMITATIONS-FUEL SYSTEM LIMITATIONS Airworthiness limitation items (ALI) are mandatory maintenance items for the fuel system that can include CDCCLs, inspections, or other procedures necessary to make sure the fuel tank flammability exposure does not increase above the certification limits as a result of maintenance actions, repairs, or alterations throughout the operational life of the airplane. The Airworthiness Limitations (AWL) gives the mandatory scheduled maintenance requirements applicable to the C Series aircraft. The fuel system limitation (FSL) information is found in the AWL section 7. This section contains task data about the aircraft fuel system maintenance tasks generated by the fuel tank system safety and ignition prevention analysis accomplished in compliance with federal aviation administration (FAA) Special Federal Aviation Regulation (SFAR) No. 88 - fuel tank system fault tolerance evaluation requirements. FSL tasks specified are mandatory airworthiness limitations and must be incorporated into the operators locally approved maintenance schedule.
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CS1_CS3_2800_005
28-00 Fuel Safety
Figure 2: Airworthiness Limitations-Fuel System Limitations Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-00 Fuel Safety
CRITICAL DESIGN CONFIGURATION CONTROL LIMITATIONS Critical design configuration control limitations (CDCCLs), a type of airworthiness limitation, include those features of the design that must be present and maintained to achieve the safety level intended by FAR 25.981 for the operational life of the airplane. The information is found in the airworthiness limitations (AWL) section 7. This section contains information about the critical design configuration control limitations (CDCCLs) generated by the fuel tank system safety and ignition prevention analysis accomplished in compliance with FAA SFAR 88 - fuel tank system fault tolerance evaluation requirements. The critical design configuration control limitations (CDCCLs) are established to preserve the critical ignition source prevention features of the fuel system design that are necessary to prevent the occurrence of an unsafe condition identified by the fuel tank safety assessment (FTSA) analysis. The purpose of the CDCCLs is to provide instructions to aircraft maintenance personnel to retain the critical ignition source prevention features during the time of configuration change that may be caused by alterations, repairs or maintenance actions. Identification of the CDCCLs design features, parts or components is provided within the applicable task procedures of the instructions for continued airworthiness (ICA), including the Aircraft Maintenance Publication (AMP). The critical ignition source prevention features of the parts or components identified in the list of CDCCLs shown in Table 2 must be maintained in a configuration identical to the approved type design for the C Series aircraft.
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CS1_CS3_2800_006
28-00 Fuel Safety
Figure 3: Critical Design Configuration Control Limitation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-00 Fuel Safety
CDCCL WARNING Identification of the CDCCL design features, parts, or components is provided within the applicable task procedures of the instructions for continued airworthiness (ICA), including the Aircraft maintenance Publication (AMP). The following warning is included when a task can impact CDCCL features.
WARNING DO NOT DAMAGE OR MODIFY THE FUEL TANK PLUMBING LINES, SELF-BONDING COUPLINGS, CONDUCTIVE FITTINGS, AND METAL-TO-METAL INTERFACE ELECTRICAL BONDING OF THE FUEL TANK COMPONENTS. MAKE SURE THAT PROTECTIVE SEALANT IS RE-APPLIED AND/OR BONDING CHECK IS DONE AS INSTRUCTED. THESE DESIGN FEATURES ARE CLASSIFIED AS CRITICAL DESIGN CONFIGURATION CONTROL LIMITATION (CDCCL) ITEMS. THEY GIVE ELECTROSTATIC AND LIGHTNING PROTECTION. DAMAGE OR MODIFICATION TO THESE ITEMS CAN POSSIBLY CAUSE AN IGNITION SOURCE INTO THE FUEL TANK BECAUSE OF ELECTROSTATICS OR A LIGHTNING STRIKE.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-8
CS130_28_L2_FuelSafety_CCDLWarning.fm
28 - Fuel
CS1_CS3_2800_007
28-00 Fuel Safety
Figure 4: Critical Design Configuration Control Limitation Warning in Aircraft Maintenance Publication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-9
CS130_28_L2_FuelSafety_CCDLWarning.fm
28 - Fuel 28-00 Fuel Safety
FUEL TANK ACCESS AND VENTING FUEL TANK ACCESS
The following is a summary of steps to remove the fuel fumes from the fuel tanks using the fuel tank venting kit:
The following is a summary of steps to prepare the fuel tank for access:
1. Put an air compressor more than 30 m (100 ft) from the aircraft.
1. Disconnect the aircraft batteries.
2. Connect the air compressor to the same ground point to which the aircraft is connected.
2. Disconnect external power and put warning placards in the flight deck ‘Do not apply electrical power to the aircraft’. 3. Establish a safety zone around the aircraft and install warning signs ‘Open fuel tank around the aircraft’. 4. After the tanks are defueled, drain the remaining fuel through the water drain valves. 5. Remove the fuel fumes from the fuel tank using the fuel tank venting kit.
3. Install the inlet wing adapter at the outboard access panel opening. 4. Connect air hose from the air compressor to the inlet wing adapter. 5. Install the outlet wing adapter at the inboard access panel opening. 6. Connect the exhaust hose to the outlet wing adapter and to a ground point. Put the exhaust hose outlet so it discharges outside of the hangar.
6. After 30 minutes, monitor the fuel tank atmosphere using a gas detector. 7. Do not enter the fuel tank until the fuel fume concentration is less than the lower explosion limit and the oxygen level is at least 19.5%. 8. Continue to vent the fuel tank when work is being done inside the fuel tank. 9. When working in the fuel tank, measure the fuel fume concentration in and around the fuel tank every 30 minutes. 10. Only personnel trained to enter a fuel tank are allowed to enter the fuel tank. 11. Remove any residual fuel using cotton cloths or sponges.
WARNING 1. MAKE SURE NO PERSON ENTERS THE FUEL TANK IF THE FUME CONCENTRATION IS MORE THAN THE LOWER EXPLOSION LIMIT. 2. MAKE SURE THAT ALL EQUIPMENT IS CLEAR OF THE EXHAUST DUCT. STATIC ELECTRICITY GENERATED BY AIRFLOW FROM THE EXHAUST DUCT CAN CAUSE AN EXPLOSION. 3. OBEY ALL LOCAL REGULATIONS WHEN DISCARDING FUEL.
FUEL TANK VENTING The fuel tank venting kit has the necessary equipment to ventilate the fuel tank including air hoses, air movers, antistatic ducts, couplings, and wing adapters. Complete instructions are available with the kit.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-10
CS130_28_L2_FuelSafety_FuelTankAccess.fm
28 - Fuel 28-00 Fuel Safety
Wing Adaptors
Wing Venting Kit
Couplings
100 mm Antistatic Ducts
Small Air Mover
200 mm to 100 mm Coupling 200 mm Antistatic Ducts
CS1_CS3_2800_008
Large Air Mover
Air Hoses
Figure 5: Fuel Tank Access and Venting Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-11
CS130_28_L2_FuelSafety_FuelTankAccess.fm
28 - Fuel 28-00 Fuel Safety
FUEL TANK ENTRY The following is a summary of steps to enter the fuel tank:
PREPARATION The following is a summary of steps to prepare for fuel tank entry: 1. Obey all local safety regulations during fuel system maintenance. 2. Firefighting extinguishers must be positioned near the aircraft. 3. Connect the grounding cable to an approved ground and to an aircraft ground point. 4. Ground the maintenance stands and equipment to the same ground point the aircraft is connected to. 5. Do not connect or disconnect electrical equipment to or from energized outlets that are less than 30.5 m (100 ft) from an open fuel tank . 6. Use explosion-proof floodlights and extension cords for external lighting. Use explosion-proof flashlights or explosion-proof lights inside the fuel tank. Explosion-proof lights must be connected to the electrical power before being put it in the fuel tank.
NOTE Only personnel with fuel tank entry training should enter a fuel tank. The local authorities may require the personnel to have a tank entry permit
1. Wear clean lint-free cotton coveralls with buttons or zippers that do not cause sparks in the fuel tanks. Do not use wool, silk, nylon, or other synthetic clothing. When entering the fuel tank, use cotton hair and shoe covers as well. 2. When entering a fuel tank, an observer must be stationed near the fuel tank opening. The observer must remain in contact with personnel inside the tank and monitor their condition. 3. Metal boxes or containers with sharp or rough edges can cause damage to the fuel tank. Use containers with rounded corners that do not generate static electricity. Sharp-edged tools must be put it in a container when not in use. 4. Use only sponges or cotton rags to absorb fuel. Other materials can cause a static discharge 5. Leave the fuel tank immediately if the fuel concentration exceeds the lower explosion limit or is below the minimum oxygen concentration level. 6. Keep a list of all the tools, equipment, and materials entering the fuel tank. All tools, equipment, and material must be accounted for before the tank is closed. 7. Make sure the fuel tank is thoroughly inspected for cleanliness before closing the fuel tank.
ENTRY Before entering a fuel tank, check that the breathing apparatus kit is in good working condition. The breathing apparatus kit has an instruction manual, pressure regulator and valve, air filter, air hoses, and the necessary accessories to supply air to a full or half face mask.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
WARNING DO NOT USE STEEL WOOL OF ANY TYPE IN THE FUEL TANKS. PIECES OF STEEL WOOL CAN BREAK OFF AND CAUSE DAMAGE OR AN EXPLOSION IF THEY TOUCH ELECTRICAL COMPONENTS.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-12
CS130_28_L2_FuelSafety_FuelTankEntry.fm
28 - Fuel 28-00 Fuel Safety
Breathing Tube Panel Filter and Regulator
Regulator Valve Air Hose
FIRE EXTINGUISHER
COTTON COVERALLS
Face Mask
EXPLOSION PROOF LIGHTS
BREATHING APPARATUS KIT COTTON RAGS
CS1_CS3_2800_009
Half Mask
Figure 6: Fuel Tank Entry Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-13
CS130_28_L2_FuelSafety_FuelTankEntry.fm
28 - Fuel 28-11 Fuel Storage
28-11 FUEL STORAGE GENERAL DESCRIPTION The aircraft fuel is stored in the integral center, right wing, and left wing tanks located within the wing box structure. The tanks are isolated by a common boundary formed by structural ribs, and have independent provisions for fuel gauging and refueling. Baffle ribs are installed in the wing to restrict fuel migration outboard during aircraft maneuvers. The wings and the center wing box are all constructed using carbon fiber reinforced polymer (CFRP), with the exception of the wing ribs, which are metallic.
One flapper check valve allows fuel to drain from the surge tank back to the main tank. Two fuel drain valves are installed in each of the fuel tanks at low points to provide water drainage.
Each wing tank has a main and a collector tank. The collector tank is partially sealed and is located inboard of the wing tank. A dry bay is located above each engine strut in the wing tanks. The dry bays prevent fuel from spilling onto an engine if a wing tank is ruptured. Access panels, located on the underside of the wing, provide access for the inspection, repair, and replacement of components inside the wing tanks. All ribs, including the baffle ribs, incorporate drain passages to prevent the accumulation of water and trapped fuel. Each rib has openings at the upper intersection with the wing inner-skin liner, which facilitates airflow within the tank. The interior of the tanks is chemically treated against corrosion and coated with a biocide compound. All joints and rivet areas are sealed to ensure tank seal integrity. Flapper check valves allow fuel to gravity feed from the main tanks to the collector tank to ensure that fuel is available at the fuel pumps in all aircraft attitudes. They also prevent the flow of fuel from the collector tank to the main tank during wing down and uncoordinated aircraft maneuvers. The check valve is a rubberized fabric hinged flapper.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-14
CS130_28_L2_FuelStorage_General.fm
28 - Fuel 28-11 Fuel Storage
RIB 5
RIB 0
Fuel Tank Access Panel (typical)
RIB 21 RIB 7
Fuel Tank Access Panel (typical)
RIB 21 RIB 23
RIB 7
Fuel Drain Valve
Flapper Check Valve
Surge Tank Main Tank Wing Dry Bay Collector Tank Center Tank
Flapper Check Valves (4x)
NOTE Fuel Drain Valve
Components from right wing shown. Left wing similar.
CS1_CS3_2811_007
LEGEND
Figure 7: Fuel Storage System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-15
CS130_28_L2_FuelStorage_General.fm
28 - Fuel 28-11 Fuel Storage
COMPONENT LOCATION The following components are installed in the fuel storage system: •
Flapper check valves
•
Fuel drain valves
•
Fuel tank access panels
FLAPPER CHECK VALVES Four flapper check valves are located on RIB 7, between the main and collector tanks on each wing. One flapper check valve is located on RIB 21, between the main tank and surge tank on each wing. FUEL DRAIN VALVES Fuel drain valves are installed at low points in the center, main, and collector tanks. The fuel drain valves are installed between the RIB 1 and RIB 2 of the left and right side of the center wing box. There is one fuel drain valve between RIB 5 and RIB 6, and another between RIB 7 and RIB 8 on each wing. FUEL TANK ACCESS PANELS Fuel tank access panels are installed on the underside of the wings and center tank, between each rib from RIB 0 to RIB 23.
WARNING BE CAREFUL WHEN YOU REMOVE THE ACCESS PANEL. FUEL CAN BE ON THE INNER SURFACE OF THE PANEL. FUEL CAN CAUSE INJURY TO PERSONS AND/OR DAMAGE TO EQUIPMENT.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-16
CS130_28_L2_FuelStorage_CompLocation.fm
28 - Fuel 28-11 Fuel Storage
RIB 5 RIB 0
Fuel Tank Access Panel (typical)
Fuel Tank RIB 7 Access Panel (typical) RIB 21 RIB 23
FUEL TANK ACCESS PANELS
Surge Tank Main Tank Wing Dry Bay Collector Tank Center Tank
FLAPPER CHECK VALVE
FUEL DRAIN VALVE
CS1_CS3_2811_002
LEGEND
Figure 8: Fuel Storage System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-17
CS130_28_L2_FuelStorage_CompLocation.fm
28 - Fuel 28-11 Fuel Storage
PRACTICAL ASPECTS FUEL DRAIN VALVE SERVICING The fuel drain valves are used for: •
Removing water and contaminants
•
Fuel sampling
•
Draining the remaining fuel after defueling a tank
The fuel drain valve is a simple poppet operated type device. O-ring seals on the valves may be replaced without removing the valve from the tank. The valve has a push-to-drain operation and can be locked in the open position using a standard screwdriver.
NOTE The fuel drain valve body is plastic. Use caution when attempting to operate in cold weather.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-18
CS130_28_L2_FuelStorage_PracticalAspects.fm
28 - Fuel 28-11 Fuel Storage
SERVICE POSITION FOR O-RING REPLACEMENT CS1_CS3_2811_006
OPEN FOR DRAINING
Figure 9: Fuel Drain Valve Servicing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-19
CS130_28_L2_FuelStorage_PracticalAspects.fm
28 - Fuel 28-12 Vent System
28-12 VENT SYSTEM GENERAL DESCRIPTION The fuel vent system maintains the pressure of the fuel tanks near the pressure of the outside atmosphere to prevent damage to the wing structure. The surge tanks are open to outside air through the vent scoops. During flight, the vent scoops maintain positive pressure inside the surge tanks. Each fuel tank is connected to a wing surge tank through a vent tube. The vent tubes have bifurcated openings at the two ends to provide added protection against blockage. The right wing tank vents through the right surge tank, and the left wing tank and the center tank vent through the left surge tank. The vent tubes keep the pressure of all the fuel tanks near the pressure in the surge tanks. Float vent valves and drains ensure the vent tubes are not blocked by fuel that might enter the lines. They also drain any fuel that may have accumulated in the vent tubes. The float drain valves close when the fuel level reaches the vent tubes. The center tank float drain valves also provide additional venting redundancy for the center tank. Center tank float vent valves are installed in the center tank to ensure venting in case the vent tube openings are submerged in fuel during aircraft maneuvers. The main tank float vent valve remains above the fuel level in any ground condition due to the aircraft wing dihedral. Any fuel entering the surge tank through the float vent valve drains back to the main tank through the surge tank flapper check valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The flow of air between the surge tanks and the atmosphere is normally through the NACA inlet scoop and flame arrestor. The inlet scoop creates a small pressure within the surge tank during flight. The inlet scoop also provides a path for fuel to drain overboard should the surge tank overflow with fuel. The flame arrestor ensures that a flame does not enter the fuel tanks through the vent system. In case of a blockage, the flame arrestor has a built-in negative pressure-relief valve to provide backup pressure equalization. In normal operation, air flows between the main wing tanks and surge tanks through the bifurcated vent openings or through the float vent valves, depending on the aircraft attitude and the level of fuel in the tank. For the center tank vent, air flows through the vent tubes, either by the bifurcated opening or through one of the float vent valves installed in the vent tube. Main wing tank pressure-relief valves allow pressure relief to the atmosphere. Center tank positive pressure is relieved into the right wing tank and negative pressure is relieved into the right surge tank. Surge tank flapper valves also provide redundant negative pressure relief to the main tanks. If the flow path is submerged by fuel and the alternate path is blocked, the internal pressure of the tank will cause fuel to be spilled into the surge tank. The fuel either drains back through the flapper valves or spills into the atmosphere through the NACA inlet scoop and flame arrestor. In case of a complete blockage in the normal flow path, the appropriate pressure-relief valve (PRV) will provide backup pressure equalization.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-20
CS130_28_L2_VentSystem_General.fm
28 - Fuel 28-12 Vent System
Center Tank Vent Tube
Left Main Tank Vent Tube
Center Tank Positive Pressure-Relief Tube
L SURGE TANK
CENTER TANK
L MAIN TANK
Right MainTank Vent Tube
R DRY BAY
L DRY BAY
Bifurcated Opening
Center Tank Negative Pressure-Relief Tube
COLLECTOR TANK
R MAIN TANK COLLECTOR TANK
R SURGE TANK
Negative Pressure-Relief Valve
LEGEND
CS1_CS3_2812_002
NACA Inlet Scoop Flame Arrestor Float Vent Valve Pressure-Relief Valve Float Drain Valve Flapper Check Valve
Figure 10: Fuel Vent System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-21
CS130_28_L2_VentSystem_General.fm
28 - Fuel 28-12 Vent System
COMPONENT LOCATION The following components are installed in the fuel vent system: •
Float drain valves
•
Center tank float vent valves
•
Wing tank float vent valves (refer to figure 12)
•
Bifurcated vent openings (refer to figure 12)
•
NACA inlet scoop (refer to figure 13)
•
Vent relief flame arrestor (refer to figure 13)
•
Positive pressure-relief valves (refer to figure 14)
•
Center tank negative pressure-relief valve (refer to figure 14)
FLOAT DRAIN VALVES Float drain valves are installed along the vent tubes in the center and wing tanks at RIB 2 and RIB 9. CENTER TANK FLOAT VENT VALVES Two float vent valves are installed in the center tank between RIB 0 and RIB 1 and at RIB 6 of the left wing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-22
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 0 RIB 6 RIB 9
FLOAT DRAIN VALVE
CENTER TANK FLOAT VENT VALVE
CS1_CS3_2812_006
RIB 9
Figure 11: Float Drain Valve and Center Tank Float Vent Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-23
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
WING TANK FLOAT VENT VALVES Wing tank float vent valves, installed near the top of RIB 21, vent each main tank into the surge tank. BIFURCATED VENT OPENINGS The bifurcated vent openings are located in the surge tanks at RIB 22, RIB 8, and at RIB 6R in the right wing only.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-24
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 6 RIB 8 RIB 8
BIFURCATED VENT OPENINGS
WING TANK FLOAT VENT VALVE
CS1_CS3_2812_007
RIB 22
RIB 22
Figure 12: Bifurcated Vent Openings and Wing Tank Float Vent Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-25
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
NACA INLET SCOOP A NACA inlet scoop is installed in each surge tank outboard of RIB 22. VENT RELIEF FLAME ARRESTOR The vent relief flame arrestor is installed in the left and the right wings between RIB 22 and RIB 23.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-26
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 22 RIB 23
Flame Arrester
NACA VENT SCOOP AND FLAME ARRESTOR
CS1_CS3_2812_009
Negative Pressure-Relief Valve
Figure 13: NACA Inlet Scoop and Vent Relief Flame Arrestor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-27
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
POSITIVE PRESSURE-RELIEF VALVES Positive pressure-relief valves are installed at the access panel on the left and the right wings between RIB 20 and RIB 21. CENTER TANK POSITIVE PRESSURE-RELIEF VALVE The positive pressure-relief valve for the center tank is installed on RIB 7 in the right wing tank. CENTER TANK NEGATIVE PRESSURE-RELIEF VALVE The center tank negative pressure-relief valve is installed on the right wing at RIB 21.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-28
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-12 Vent System
RIB 7
RIB 21
RIB 21
CENTER TANK POSITIVE PRESSURE-RELIEF VALVE
CENTER TANK NEGATIVE PRESSURE-RELIEF VALVE
CS1_CS3_2812_008
POSITIVE PRESSURE-RELIEF VALVE
Figure 14: Positive Pressure-Relief Valves, Center Tank Positive, and Center Tank Negative Pressure-Relief Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-29
CS130_28_L2_VentSystem_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
28-23 REFUEL/DEFUEL SYSTEM GENERAL DESCRIPTION Aircraft refueling and defueling is carried out using a single-point refuel/defuel adapter, located at the right wing. The refuel adapter connects to the refuel manifold that runs near the rear spar of the fuel tanks. The refuel/defuel tube connects the refuel adapter to all of the tanks via the refuel shutoff valves.
The refuel adapter must be set to the defuel position before connecting the fuel truck hose. Defueling the aircraft can only be done in MANUAL mode.
Each tank has one refuel shutoff valve and they connect the refuel/defuel tube to the tanks. The refuel control solenoid valve controls the opening and closing of the refuel shutoff valve. The refuel control solenoid valve is controlled by the fuel quantity computer (FQC). When the solenoid is energized, fuel pressure is directed to open the refuel shutoff valve. When the refuel shutoff valve opens, fuel discharges into the tank through a diffuser. Refueling and defueling is controlled from the refuel/defuel control panel (RDCP) located on the fuselage at right wing root. Pressure refueling of the aircraft is carried out in manual or automatic mode, through the appropriate AUTO/MANUAL switch selection on the RDCP. The recommended fuel pressure for refueling is 50 psig. The FQC receives power from DC ESS BUS 1 and DC ESS BUS 2 during normal operation. The refuel shutoff valves receive power from DC ESS BUS 1 during normal operation. When the RDCP door panel is opened, the power for the FQC and the refuel shutoff valves is switched to the DC EMER BUS. The RDCP is powered by the FQC through the door switch. A second contact on the door switch provides a signal that the RDCP door is open. The aircraft is defueled by connecting the refueling manifold to the engine feed manifold. The defuel/isolation transfer shutoff valve connects the manifolds. The valve is powered and controlled by the FQC. When DEFUEL is selected on the RDCP, the valve opens and DEFUEL is displayed in the PRESEL window. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-30
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
Fuel Door Switch FUEL QUANTITY COMPUTER REFUEL / DEFUEL PANEL ON
L ENG
Defuel/Isolation Transfer Shutoff Valve
CENTER
RIGHT
1ØØØ
ØØØØ
2ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
Left Wing Refuel Solenoid Shutoff Valve
LEFT
LEFT
CENTER
RIGHT OPEN
DEFUEL
AUTO
DC ESS BUS 2 DECR
REFUEL
DC EMER BUS
Right Wing Refuel Solenoid Shutoff Valve
R ENG
START
CLOSE
MANUAL
DC ESS BUS 1
28 VDC
PRESEL INCR
MANUAL STOP/SOV TEST
Refuel/Defuel Adapter
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
L SURGE TANK
R MAIN TANK
COLLECTOR TANK
COLLECTOR TANK
R SURGE TANK
SOL
M
Left Wing Refuel Shutoff Valve
Manual Defuel Valve
LEGEND Center Tank Refuel Shutoff Valve Manual Defuel Center Tank Refuel Right Wing Refuel Valve Lever Solenoid Shutoff Valve Shutoff Valve
M PS SOL
Motor Shutoff Valve Pressure Switch Solenoid Shutoff Valve
CS1_CS3_2823_027
M
Figure 15: Refuel/Defuel System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-31
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUELING Auto Refueling With the AUTO/MANUAL switch in AUTO, the fuel quantity to be delivered is preselected on the RDCP using the PRESEL INCR/DECR switch setting. The REFUEL START/STOP/SOV TEST switch has a momentary STOP/TEST SOV, a START position, and a center OFF position. The START position is used for automatic refueling. When refueling is in progress in AUTO MODE, selecting STOP/SOV TEST stops refueling. The refueling operation starts on selection of the REFUEL/DEFUEL switch to REFUEL, and the START/STOP/SOV TEST switch to the START position. The refueling operation only stops when the preselected fuel quantity is delivered, or the operation is stopped by selecting the START/STOP/SOV TEST switch to the STOP/SOV TEST position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-32
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ
2ØØØ
TOTAL
PRESEL
OFF
Left Wing Refuel Solenoid Shutoff Valve
LEFT
L ENG
CENTER
RIGHT OPEN
DEFUEL
AUTO
PRESEL INCR
DECR
START
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Refuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
R SURGE TANK
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
SOL
M
Left Wing Refuel Shutoff Valve
Center Tank Refuel Shutoff Valve APU Manual Defuel Center Tank Refuel Right Wing Refuel Valve Lever Solenoid Shutoff Valve Shutoff Valve
CS1_CS3_2823_003
AC
Figure 16: Auto Refueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-33
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
Manual Refueling The manual mode is entered by selecting the MANUAL on the RDCP and the REFUEL/DEFUEL switch to the REFUEL position. MANUAL is displayed in the preselect window. The refuel SOVs are opened by setting the MANUAL REFUEL VALVE switches to OPEN. The individual tank quantities are monitored on the RDCP. When the desired tank quantity is reached, the respective refuel SOVs are closed manually using the MANUAL REFUEL VALVE switches on the RDCP. In the manual mode, refueling continues until the MANUAL REFUEL VALVE switches are closed, or the fuel high-level condition is reached. At that point, the refuel valves are closed by the FQC. These switches are hardwired to the FQC to provide a discrete to open or close their respective valve in case the RDCP fails.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-34
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ OFF
Left Wing Refuel Solenoid Shutoff Valve
MANUAL PRESEL
TOTAL LEFT
L ENG
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
DECR
START
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Refuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
R SURGE TANK
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
SOL
M
Left Wing Refuel Shutoff Valve
Center Tank Refuel Shutoff Valve APU Manual Defuel Center Tank Refuel Right Wing Refuel Valve Lever Solenoid Shutoff Valve Shutoff Valve
CS1_CS3_2823_004
AC
Figure 17: Manual Refueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-35
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
DEFUELING Suction Defueling Suction defueling is carried out using the fuel truck suction pressure. With the RDCP and refuel adapter configured for defueling, the fuel truck pumps are turned on. The maximum suction pressure allowed is -8 psig. Fuel from the left and right main tank is suctioned out through the inlet snorkel of the AC boost pumps. Fuel in the center tank is removed by opening the manual defuel valve on the rear spar of the center tank. Fuel is drawn out of the center tank through the defuel valve using fuel truck suction pressure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-36
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ OFF
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Defuel/Isolation Transfer Shutoff Valve
DEFUEL PRESEL
TOTAL LEFT
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
DECR
START
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Defuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
L SURGE TANK
COLLECTOR TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
R SURGE TANK
SOL
M
Manual Defuel Valve
APU
Manual Defuel Center Tank Refuel Valve Lever Solenoid Shutoff Valve
CS1_CS3_2823_014
AC
Figure 18: Suction Defueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-37
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
Pressure Defueling Pressure defueling is carried out using the aircraft AC boost pumps. With the RDCP and refuel adapter configured for defueling, the AC boost pumps are turned on manually. The fuel is pumped out of the main tanks through the engine feed manifold to the refuel manifold. If there is fuel in the center fuel tank, the AC boost pumps provide the motive flow to operate the transfer pumps. The transfer pumps empty the center tank fuel into the main tanks, where it is then pumped out.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-38
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
ON
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ OFF
FUEL PANEL
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Defuel/Isolation Transfer Shutoff Valve
DEFUEL PRESEL
TOTAL LEFT
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
DECR
START
Engine Feed Manifold
Right Wing Refuel Solenoid Shutoff Valve
R ENG
CLOSE REFUEL
MANUAL
MANUAL STOP/SOV TEST
SOL
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
R MAIN TANK
LEGEND Defuel AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
L SURGE TANK
COLLECTOR TANK
COLLECTOR TANK
M
Motor Shutoff Valve
M
R SURGE TANK
One-Way Check Valve PS
SOL
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
SOL
M
APU
Manual Defuel Center Tank Refuel Valve Lever Solenoid Shutoff Valve
Refuel Manifold
CS1_CS3_2823_005
AC
Figure 19: Pressure Defueling Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-39
CS130_28_L2_RefuelDefuelSyst_General.fm
28 - Fuel 28-23 Refuel/Defuel System
COMPONENT LOCATION The following components are installed in the pressure refuel/defuel system: •
Refuel adapter
•
Refuel/defuel control panel
•
Refuel shutoff valve
•
Refuel control solenoid valve
•
Defuel/isolation transfer shutoff valve and actuator
•
Manual defuel valve and lever
REFUEL ADAPTER The adapter is located on the right wing outboard of the engine at RIB 13R. An optional refuel adapter is available and installed at RIB 13L. REFUEL/DEFUEL CONTROL PANEL The refuel/defuel control panel (RDCP) is located on the fuselage near the inboard right wing leading edge.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-40
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
Refuel/Defuel Control Panel
CS1_CS3_2823_026
Refuel Adapter
Figure 20: Refuel Adapter and Refuel/Defuel Control Panel Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-41
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL SHUTOFF VALVE The refuel shutoff valves for each main tank are mounted on the inboard side of RIB 12 inside the fuel tank. The center tank refuel shutoff valve is mounted on the inboard side of RIB 1L inside the center fuel tank. REFUEL CONTROL SOLENOID VALVE The main tank refuel control solenoid valves are mounted on a bracket at RIB 12 on the front spar. The center tank refuel control solenoid valve is mounted on the aft spar at RIB 1L.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-42
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
RIB 1L
RIB 12R
RIB 12L
Front Spar RIB 12L
Refuel Control Solenoid Valve
Refuel Control Solenoid Valve
CENTER WING BOX
LEFT WING SHOWN (RIGHT WING SIMILAR)
CS1_CS3_2823_006
Refuel Shutoff Valve
Figure 21: Refuel Control Solenoid Valve and Refuel Shutoff Valve Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-43
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
DEFUEL/ISOLATION TRANSFER SHUTOFF VALVE AND ACTUATOR The valve body is mounted inside the center tank between RIB 0 and RIB 1L on the rear spar. The 28 VDC motor actuator is mounted outside of the tank. MANUAL DEFUEL SHUTOFF VALVE AND LEVER The center tank manual defuel shutoff valve is mounted inside the center tank on the rear spar between RIB 1L and RIB 2L. A manual lever, mounted on the outside of the tank, is used to operate the valve. The valve lever is pushed in to move it open or closed. The lever assembly and valve body can be replaced independent of one another.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-44
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
Defuel/Isolation Transfer Valve
B A
RIB 0
RIB 1L
Defuel/Isolation Transfer Valve Actuator
A MANUAL DEFUEL VALVE AND LEVER
CS1_CS3_2823_025
B
Figure 22: Defuel System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-45
CS130_28_L2_RefuelDefuelSyst_CompLocation.fm
28 - Fuel 28-23 Refuel/Defuel System
COMPONENT INFORMATION REFUEL ADAPTER The refuel adapter provides a connection point for single-point defueling/refueling of the aircraft. The refuel adapter has a poppet assembly and a fuel cap attached to the adapter by a cable assembly. During defueling, a locking mechanism is rotated to the DEFUEL position, allowing fuel to flow from the refuel manifold to the truck.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-46
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel 28-23 Refuel/Defuel System
Refuel/Defuel Selector
Refuel Adapter
CS1_CS3_2823_008
FUEL CAP REMOVED
Figure 23: Refuel Adapter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-47
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL/DEFUEL CONTROL PANEL The RDCP controls and monitors aircraft refueling and defueling. The RDCP functions are enabled when the aircraft is on the ground and the RDCP refueling door is in the open position. The refuel door switch signals the fuel quantity computer (FQC) that the refuel door is open.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-48
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel
Refuel Door Switch
Refuel Door Target
CS1_CS3_2823_011
28-23 Refuel/Defuel System
Figure 24: Refuel/Defuel Control Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-49
CS130_28_L2_RefuelDefuelSyst_CompInfo.fm
28 - Fuel 28-23 Refuel/Defuel System
CONTROLS AND INDICATIONS REFUEL/DEFUEL CONTROL PANEL The refuel/defuel control panel (RDCP) controls and monitors the aircraft refueling and defueling operations. An indicator displays the left, center, and right fuel tank quantities, the total fuel quantity, and the load preselect quantity. The panel has switches to control power, automatic, or manual modes, and refuel and defuel modes.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-50
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ
5ØØØ
TOTAL
PRESEL
OFF
LEFT
CENTER
RIGHT OPEN
DEFUEL
AUTO
PRESEL INCR
DECR
START
CLOSE REFUEL
MANUAL STOP/SOV TEST CS1_CS3_2823_023
MANUAL
Figure 25: Refuel/Defuel Control Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-51
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
RDCP Messages
RS-422 COMMUNICATION ERROR
Messages related to refueling and defueling display on the RDCP.
A fault has occurred in the communications bus and the RDCP is inoperative. The displays remain dashed and the PRESEL portion of the window COMM ERR is displayed. The RDCP does not respond to any switch inputs.
INHIBIT When refueling the aircraft in manual mode or defueling the aircraft, an INHIBIT message is displayed whenever the RCDP is powered in manual mode or when switched to manual mode and any of the MANUAL LEFT/CENTER/RIGHT switches are in the OPEN position. A CLOSE ALL VALVES message is also displayed on the panel. The message clears when all of the MANUAL LEFT/CENTER/RIGHT switches are closed.
RCDP SWITCH FAULT The RDCP detects a switch fault (a momentary switch is held in the same position for more than 3 minutes). The message SWCH FLT will be displayed on the preselect window. The fuel tank quantities will continue to be displayed on the RDCP.
FULL
RDCP FAULT
A FULL message indicates the level in the respective tank is between 97% and 98% of total capacity. The message remains as long as the condition exists.
For other RDCP faults, the message RDCP FLT is displayed on the preselect window. If possible, the RDCP continues to display the fuel tank quantities. The quantities displayed must be verified with information from the flight deck.
HIGH LVL A HIGH LVL message indicates the level in the respective tank is greater than 98% of total capacity. The message remains as long as the condition exists.
FQC FAULT An FQC fault that affects the refueling process is detected, the message FQC FLT is displayed on the preselect window. If possible, the RDCP continues to display the quantities.
IMBAL During refueling, an IMBAL message is displayed, flashing alternately with the quantity information when there is an imbalance between the left and right tank fuel quantities. A fuel quantity difference of 364 kg (800 lb) or greater between the wing fuel tanks is considered an imbalance. An IMBAL message is displayed for the fuel tank containing the lesser of the two imbalance quantities.
AUTOMATIC REFUELING STOPPED An automatic refuel has been stopped before completion (START/STOP switch toggled to the STOP position during an automatic refuel), the STOPPED message is displayed in the preselect window for 5 seconds.
RCDP CPU FAULT A fault has occurred in the RDCP CPU. The displays remain blank and the RDCP does not respond to any switch inputs. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-52
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
CLOSE
ALL
4ØØØ
RIGHT
REFUEL / DEFUEL PANEL LEFT
CENTER
VALVES
FULL
INHIBIT
61ØØ
TOTAL
PRESEL
2000
HIGH LVL
PRESEL
CENTER
CENTER
RIGHT
1000
CENTER
RIGHT
TOTAL
PRESEL
REFUEL / DEFUEL PANEL LEFT
CENTER
----
----
4ØØØ
10000
----
----
TOTAL
PRESEL
PRESEL
CPU FAULT
FUEL TANK IMBALANCE
REFUEL / DEFUEL PANEL LEFT
LEFT
2000
2000
5ØØØ
SWC H F LT
TOTAL
PRESEL
1000
TOTAL
CENTER
---COMM ERR PRESEL
REFUEL / DEFUEL PANEL
RIGHT
LEFT
2000
2000
5ØØØ
R D C P F LT
5ØØØ
TOTAL
PRESEL
TOTAL
SWITCH FAULT MESSAGE
RIGHT
RS-422 COMMUNICATION ERROR
REFUEL / DEFUEL PANEL
RIGHT
2000
CENTER
2000 10000
2000 TOTAL
RIGHT
FUEL TANK HIGH LEVEL
REFUEL / DEFUEL PANEL LEFT
1000
61ØØ
FUEL TANK FULL
REFUEL / DEFUEL PANEL
IMBAL
LEFT
10000
TOTAL
INHIBIT
LEFT
1000
REFUEL / DEFUEL PANEL
RIGHT
1000
CENTER
1000
RIGHT
2000 F Q C F LT PRESEL
FQC FAULT
RDCP FAULT
REFUEL / DEFUEL PANEL CENTER
2000
RIGHT
1000
2000
5ØØØ
STOPPED
TOTAL
PRESEL
AUTOMATIC REFUELING STOPPED
CS1_CS3_2823_017
LEFT
Figure 26: Refuel/Defuel Control Panel Messages Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-53
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
OPTIONAL VIRTUAL REFUEL PANEL The virtual refuel panel is displayed on the FUEL synoptic page. The virtual refuel panel is displayed when the aircraft is on ground only with no engines running. To refuel the aircraft using the virtual refuel panel, the refuel door must be open, the RDCP powered, and the REFUEL switch set to AUTO. This panel allows the crew to enter the total fuel quantity in the preselect data entry field. Pressing ENTER sends the data to the FQC. If the preselect value entered exceeds the tank capacity, maximum allowable fuel quantity is assumed by the FQC. If refueling is in progress, the preselect fuel quantity can be modified. If the preselect value is less than the current fuel load, refueling stops at the current fuel quantity level. When the virtual refuel panel is in control of the preselect quantity, the RDCP ignores inputs on the RDCP PRESEL INCR/DECR switch. The refuel SOVs cannot be opened from the flight deck. The START switch on the RDCP must be selected to open the aircraft refuel SOVs. The virtual refuel panel has priority over the RDCP commands. If no value or 0 is entered in the preselect field of the virtual refuel panel, the preselect fuel quantity can be entered in the RDCP. Cycling the RDCP power for 10 seconds overrides the flight deck preselect field. The source of the preselect fuel quantity used by the FQC is displayed on the virtual refuel panel and on the RDCP. The left, center, right, and total fuel quantities are still displayed on the RDCP in their respective windows. The PRESEL window displays the message COCKPIT alternating with the preselect quantity that is entered in the flight deck refuel panel. The COCKPIT message indicates that the entry has been made in the flight deck and that refueling may be started.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-54
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
3ØØØ
43 Ø Ø Ø
3ØØØ
ON
10000 KG 0 KG
1 Ø3 Ø Ø Ø OFF
COCKPIT DECR
PRESEL
TOTAL LEFT
PRESEL INCR
CENTER
RIGHT OPEN
DEFUEL
AUTO
START
CLOSE
MANUAL REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
REFUEL
ACCEPTED
MANUAL STOP/SOV TEST
VALUE SET TO MAX
34000 KG IN PROGRESS
REFUEL MANAGED BY COCKPIT - COMPLETE
20 °C 3000 KG
20 °C 3000 KG
REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED 34000 KG COMPLETE
REFUEL MANAGED BY PANEL - COMPLETE
4000 KG
REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED èèèèè KG STANDBY
APU
REFUEL PRESELECT REFUEL PRESELECT
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED 15000 KG
CKPT
EXT
REQUESTED 15000 KG
ACCEPTED èèèèè KG COMPLETE
ACCEPTED VALUE INVALID
CS1_CS3_2823_009
REQUESTED VALUE ENTERED - SYSTEM NOT READY
Figure 27: Virtual Refuel Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-55
CS130_28_L2_RefuelDefuelSyst_ControlsIndications.fm
28 - Fuel 28-23 Refuel/Defuel System
OPERATION NOTE
AUTO REFUEL OPERATION
1. If a fuel imbalance greater than 364 kg (800 lb) occurs between the wing fuel tanks, refueling must be stopped.
1. Open the REFUEL panel door. 2. Set the ON/OFF panel switch to the ON position. 3. Set the AUTO/MANUAL switch to the AUTO position. 4. Hold the PRESEL INCR/DECR switch to the INCR position to increase the desired refuel quantity displayed in the PRESEL window, or hold the switch to the DECR position to decrease the desired refuel quantity displayed in the PRESEL window.
2. When all tanks are refueled simultaneously; the time to completely refuel the center tank is approximately 30 minutes and 20 minutes to completely refuel the wing tanks. 3. If the refueling operation is stopped by selecting the START/STOP/SOV TEST to STOP/SOV TEST, a STOPPED message is displayed for 5 seconds.
5. Perform the refuel valve SOV TEST. The STOP/SOV TEST position is used to test the SOV operation at the start of refueling. The refuel valves are tested by applying fuel pressure and then selecting the SOV TEST position. An SOV TEST IN PROGRESS message appears in the display windows, alternating every 2 seconds with the fuel quantity display. The fuel valves close and an SOV TEST PASSED message is displayed on the RDCP indicating a successful test. The message is removed when any RDCP switch is selected. 6. Set the START/STOP/SOV TEST switch to the START position. The PRESEL window then displays a clock icon and the preselected fuel quantity message. 7. The refuel flow to the aircraft stops automatically when the preselected quantity is achieved. In the PRESEL window COMPLETE is displayed, alternating every 2 seconds with the preselected fuel quantity. 8. Set the ON/OFF panel switch to the OFF position and close the REFUEL panel door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-56
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL ON
LEFT
CENTER
1ØØØ
ØØØØ
REFUEL / DEFUEL PANEL
RIGHT
1ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL RESEL
OFF
LEFT
CENTER
RIGHT
DEFUEL
AUTO
PRESEL INCR
LEFT
CENTER
PASS
PASS
ON
SOV TEST
TOTAL
PRESEL
LEFT
START
PASS
2ØØØ OFF
DECR
RIGHT
CENTER
RIGHT OPEN
DEFUEL
AUTO
PRESEL INCR
DECR
START
CLOSE MANUAL STOP/SOV TO T TEST
MANUAL
REFUEL / DEFUEL PANEL LEFT
CENTER
1ØØØ
ØØØØ
REFUEL / DEFUEL PANEL
RIGHT
1ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
OPEN
DEFUEL
AUTO
PRESEL INCR
LEFT
CENTER
RIGHT
2ØØØ
ØØØØ
2ØØØ
ON
4ØØØ
5ØØØ
TOTAL
PRESEL
OFF
DECR
LEFT
START
CENTER
SOV TEST
in
MANUAL STOP/SOV MANUA TEST
RIGHT
PROGRESS
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
LEFT
CENTER
RIGHT OPEN
DECR
CLOSE REFUEL
DEFUEL
AUTO
REFUEL
MANUAL
REFUEL / DEFUEL PANEL LEFT
ON
PRESEL INCR
START
CLOSE
MANUAL
MANUAL STOP/SOV TEST
MANUAL MANUA STOP/SOV TEST
REFUEL / DEFUEL PANEL
PRESEL INCR
LEFT
CENTER
RIGHT
25ØØ
ØØØØ
25ØØ
ON
DECR
5ØØØ
COMPLETE
TOTAL
PRESEL
OFF
LEFT
START
CENTER
RIGHT OPEN
MANUAL STOP/SOV MANUA TEST
DEFUEL
AUTO
PRESEL INCR
DECR
START
MANUAL A STOP/SOV TEST
CS1_CS3_2823_012
ON
REFUEL
MANUAL
Figure 28: Automatic Refuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-57
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
MANUAL REFUEL OPERATION 1. Open the REFUEL panel door. 2. Set the ON/OFF panel switch to the ON position. 3. Set the MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 4. Set the REFUEL/DEFUEL switch to REFUEL. 5. Set the AUTO/MANUAL switch to the MANUAL position. 6. Set the MANUAL LEFT/CENTER/RIGHT switches to the OPEN position for the tanks that are to be filled. 7. Set the MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position when the required fuel level is reached. 8. Set the ON/OFF panel switch to the OFF position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-58
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL ON
LEFT
CENTER
1ØØØ
ØØØØ
RIGHT
1ØØØ
2ØØØ
5ØØØ
TOTAL
PRESEL
OFF
LEFT
CENTER
RIGHT OPEN
AUTO
DEFUEL
PRESEL INCR
DECR
START
CLOSE E
REFUEL / DEFUEL PANEL
MANUAL STOP/SOV TEST
REFUEL
MANUAL UAL
LEFT
CENTER
3ØØØ
3ØØØ
ON
9ØØØ OFF
CENTER
PRESEL RIGHT OPEN
CENTER
2ØØØ
1ØØØ
ON
5ØØØ OFF
CENTER
RIGHT
2ØØØ PRESEL
RIGHT OPEN
MANUAL
PRESEL INCR
START
REFUEL
MANUAL STOP/SOV TEST
DEFUEL
AUTO
DECR
START
REFUEL
CS1_CS3_2823_013
CLOSE
MANUAL
AUTO
MANUAL
TOTAL LEFT
DEFUEL
DECR
CLOSE
REFUEL / DEFUEL PANEL LEFT
3ØØØ
PRESEL INCR
MANUAL
TOTAL LEFT
RIGHT
MANUAL STOP/SOV TEST
Figure 29: Manual Refuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-59
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
PRESSURE DEFUEL OPERATION 1. Connect external AC power to the aircraft. 2. Open the REFUEL panel door. 3. Set the ON/OFF switch to the ON position. 4. Set all MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 5. Set the MANUAL REFUEL/DEFUEL switch to DEFUEL. Verify that the defuel message is displayed on the RDCP window. 6. On the pressure refuel adapter, set the REFUEL/DEFUEL selector to DEFUEL. 7. Connect the refuel hose nozzle to the refuel adapter. 8. Set all MANUAL LEFT/CENTER/RIGHT switches to the OPEN position. 9. Set the GRAV XFR PBA to the CLOSED position. 10. Set the BOOST PUMP switches to the ON position for the fuel tank(s) to be defueled. 11. When an AC boost pump low-pressure indication is displayed on the FUEL synoptic page, set the BOOST PUMP switch to the OFF position. 12. Set all MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 13. Set the ON/OFF switch to the OFF position. 14. Disconnect the refuel hose nozzle from the refuel adapter and turn the pressure refuel adapter REFUEL/DEFUEL selector to DEFUEL.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-60
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
FUEL MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
AUTO
OFF
ON
AUTO
OFF
R CTR
ON
GRAV XFR
REFUEL / DEFUEL PANEL
ON
LEFT
CENTER
RIGHT
2ØØØ
2ØØØ
2ØØØ
ON
FUEL PANEL
6ØØØ Refuel/Defuel Selector
OFF
DEFUEL PRESEL
TOTAL LEFT
CENTER
PRESEL INCR
RIGHT OPEN
DEFUEL
AUTO
DECR
START
CLOSE REFUEL
MANUAL STOP/SOV TEST
CS1_CS3_2823_024
MANUAL
(FUEL CAP REMOVED)
Figure 30: Pressure Defuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-61
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
SUCTION DEFUEL OPERATION 1. Open the REFUEL panel door. 2. Set the ON/OFF switch to the OFF position. 3. Set each of the MANUAL LEFT/CENTER/RIGHT switches to the CLOSE position. 4. If the center fuel tank is to be defueled, the MANUAL DEFUEL valve is opened by pushing and turning the MANUAL DEFUEL LEVER, located on the rear spar of the center fuel tank. 5. On the pressure refuel adapter, set the REFUEL /DEFUEL selector to DEFUEL. 6. Set the REFUEL/DEFUEL switch to the DEFUEL position. 7. Set the POWER SWITCH to the ON position. The PRESEL display window shows DEFUEL. 8. Set the AUTO/MANUAL switch to the MANUAL position. 9. Attach the hose to the refuel adapter. Start the fuel truck pump, which must be configured in the suction mode. 10. Set all MANUAL LEFT/CENTER/RIGHT switches to the OPEN position. 11. Fuel is suctioned out of the wing tanks. 12. Monitor the LEFT/CENTER/RIGHT display window and stop the fuel truck pump when desired quantity is reached. 13. Manually close center tank MANUAL DEFUEL valve when the desired quantity in the center wing tank is reached. 14. Set the MANUAL LEFT/RIGHT switch(es) to the CLOSE position to stop defueling in all tanks. 15. Set the ON/OFF switch to the OFF position. 16. Disconnect the refuel hose nozzle from the refuel adapter and turn the pressure refuel adapter REFUEL/DEFUEL selector to DEFUEL. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-62
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
REFUEL / DEFUEL PANEL LEFT
CENTER
RIGHT
2ØØØ
2ØØØ
2ØØØ
ON
6ØØØ OFF
DEFUEL
CENTER
DECR
PRESEL
TOTAL LEFT
PRESEL INCR
RIGHT OPEN
DEFUEL
AUTO
START
CLOSE
MANUAL
REFUEL
MANUAL STOP/SOV TEST
CS1_CS3_2823_015
MANUAL DEFUEL VALVE
Figure 31: Suction Defuel Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-63
CS130_28_L2_RefuelDefuelSyst_Operation.fm
28 - Fuel 28-23 Refuel/Defuel System
DETAILED DESCRIPTION REFUEL POWER The fuel system has two power modes for normal and battery only refuel/defuel operation. In normal operation, the refuel panel door is closed and the system is powered by DC ESS BUS 1 and DC ESS BUS 2. When the aircraft is refueled, the refuel door is open. The refuel door switch signals the FQC that the door is opened. When the RDCP power switch is selected on, the FQC receives a 28 VDC signal indicating the refuel panel is turned on. The refuel door switch also energizes relays in CDC 1 and CDC 2 to switch the fuel power from the DC ESS BUSES to the DC EMER BUSES.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-64
CS130_28_L3_RefuelDefuelSyst_DetailedDescription.fm
28 - Fuel 28-23 Refuel/Defuel System
LEFT REFUEL SHUTOFF VALVE
CDC 1-7-21 L Fuel Qty DC ESS BUS 1
CDC 1
CENTER REFUEL SHUTOFF VALVE
RIGHT REFUEL SHUTOFF VALVE
SSPC 5A
Logic: Always On 5
Refuel Valve CTL A
Logic: Always On DC EMER BUS
5
Refuel Valve CTL B K3
REFUEL PANEL DOOR SWITCH
FUEL QUANTITY COMPUTER (FQC)
LEFT REMOTE DATA CONCENTRATOR (RDC)
CENTER REMOTE DATA CONCENTRATOR (RDC)
Lane 2
Lane 1
Lane 2
Lane 1
SSPC 5A
Lane 2
CDC 2 Door Open
R Fuel Qty DC ESS BUS 2
Channel A
CDC 2-7-20
Lane 1
K3 Channel B
DC EMER BUS
RIGHT REMOTE DATA CONCENTRATOR (RDC)
DEFUEL/ISOLATION TRANSFER SHUTOFF VALVE
REFUEL/DEFUEL CONTROL PANEL
On
Refuel Power On
CS1_CS3_2823_019
Off
Figure 32: Refuel Power (Normal Operation - Door Closed) Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-65
CS130_28_L3_RefuelDefuelSyst_DetailedDescription.fm
28 - Fuel 28-23 Refuel/Defuel System
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM The following refuel/defuel components are tested through the onboard maintenance system (OMS). Refuel Solenoid Command Test When the OMS commands a refuel solenoid command test, the selected FQC channel energizes each refuel solenoid and verifies that the solenoid is energized. The FQC channel then de-energizes each refuel solenoid and verifies that the solenoid is de-energized. This test checks the electrical command of the refuel solenoids, but does not verify fuel flow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-66
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-23 Refuel/Defuel System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-REFUEL SOLENOID COMMAND TEST
001 /004 REFUEL SOLENOID COMMAND TEST
*** CAUTION *** KEEP A/C FUEL PUMPS OFF
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2823_016
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 33: Refuel Solenoid Command Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-67
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-23 Refuel/Defuel System
Defuel Shutoff Valve Command Test When the OMS commands a defuel shutoff valve command test, the selected FQC channel commands the DEFUEL SOV open, and verifies that the SOV position feedback has reported that the valve has opened. The FQC closes the defuel SOV and verifies that the SOV position feedback reports that the valve has closed. This test checks the electrical command of the defuel shutoff valve, but does not verify fuel flow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-68
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-23 Refuel/Defuel System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-DEFUEL SHUTOFF VALVE COMMAND TEST
001 /004 DEFUEL SHUTOFF VALVE COMMAND TEST
*** CAUTION *** KEEP A/C FUEL PUMPS OFF
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2823_021
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 34: Defuel Shutoff Valve Command Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-69
CS130_28_L2_RefuelDefuelSyst_PracticalAspects.fm
28 - Fuel 28-21 Engine Feed System
28-21 ENGINE FEED SYSTEM GENERAL DESCRIPTION The feed system consists of one engine feed ejector pump and one AC boost pump in the collector tank. Each engine has its own feed system. The two engine feeds are connected to allow either AC boost pump to feed both engines. Check valves in the engine feed lines prevent a loss of fuel pressure when an ejector feed pump or boost pump fails. Normally-open engine feed shutoff valves in each engine fuel feed line stop the flow of fuel in the event of an engine fire. Each engine feed system is monitored by two pressure switches. Low fuel pressure at the engine is detected by a pressure switch in the engine feed line. Each AC boost pump has a pressure switch at the pump outlet to detect low-pressure. In the event of a total failure of the engine feed pumps, the engine can suction feed through the AC boost pump inlet up to 25,000 ft. The engine-driven pump (EDP) suctions fuel through the engine feed ejector induced port and the AC boost pump inlet snorkel. Motive flow fuel generated by the EDP powers the onside engine feed and scavenge ejector pumps. A check valve in the motive flow line isolates the engine from the fuel tank. The engine-feed ejector pump is the main source of fuel for engine operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-70
CS130_28_L2_EFS_General.fm
28 - Fuel 28-21 Engine Feed System
L ENG
R ENG
SOL
L DRY BAY
CENTER TANK
PS
SOL
R DRY BAY PS PS
AC
AC
PS
M
M M
L MAIN TANK
LEGEND
AC
COLLECTOR TANK
L SURGE TANK
Engine Feed Motive Flow AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
R SURGE TANK
M
SOL
M
One-Way Check Valve PS
Pressure Switch
Engine Feed Ejector Pump
APU
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
CS1_CS3_2821_003
Motor Shutoff Valve
M
SOL
R MAIN TANK
Figure 35: Engine Feed System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-71
CS130_28_L2_EFS_General.fm
28 - Fuel 28-21 Engine Feed System
COMPONENT LOCATION The following components are installed in the engine fuel feed system: •
Engine feed ejector pump
•
AC boost pump
•
Engine feed shutoff valve and actuator
•
Engine feed pressure switch
ENGINE FEED EJECTOR PUMP The engine feed ejector pump is located in the collector tank outboard of RIB 5. AC BOOST PUMP An AC boost pump is mounted on RIB 6 in each collector tank. ENGINE FEED SHUTOFF VALVE AND ACTUATOR The engine feed shutoff valve and actuator are installed between RIB 8 and RIB 9. ENGINE FEED PRESSURE SWITCH The engine feed pressure switch is mounted on the rear spar, near RIB 8.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-72
CS130_28_L2_EFS_CompLocation.fm
28 - Fuel 28-21 Engine Feed System
RIB 5L Engine Feed Ejector Pump Rear Spar
Engine Feed Shutoff Valve AC Boost Pump
Engine Feed Pressure Switch
Engine Feed Shutoff Valve Actuator
NOTE
CS1_CS3_2821_018
RIB 9L
AC Boost Pump Pressure Switch
Left wing shown, right wing similar.
Figure 36: Engine Feed Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-73
CS130_28_L2_EFS_CompLocation.fm
28 - Fuel 28-21 Engine Feed System
COMPONENT INFORMATION AC BOOST PUMP An AC boost pump is mounted on RIB 6 in each collector tank. The AC boost pump consists of a cannister and cartridge elements. The cannister, which contains an impeller, and suction and outlet ports, is located inside the collector tank. The suction port is connected to a pickup fitted with an inlet strainer. The pump outlet is connected to the engine feed lines. The cartridge element is mounted in the dry bay. The cartridge element has a 115 VAC, variable frequency, 3-phase motor and pressure switch. The motor drives the impeller and the pressure switch monitors the pump discharge pressure. The cannister has a built-in check valve that allows the cartridge to be removed without draining the fuel tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-74
CS130_28_L2_EFS_CompInfo.fm
28 - Fuel 28-21 Engine Feed System
RIB 5L Collector Tank RIB 5L
Pump Outlet to Engine Fuel Shutoff Valve
RIB 7L
Pump Pickup
Crossfeed Line
Dry Bay
AC Boost Pump Cannister NOTE Pressure Switch
Left wing shown, right wing similar.
CS1_CS3_2821_006
AC Boost Pump Cartridge
Figure 37: AC Boost Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-75
CS130_28_L2_EFS_CompInfo.fm
28 - Fuel 28-21 Engine Feed System
CONTROLS AND INDICATIONS ENGINE FIRE PBA The L or R ENG FIRE PBA closes the engine fuel shutoff valve. FUEL PANEL The AC L or R BOOST PUMP rotary switches are on the FUEL panel. The FUEL panel is located on the overhead panel.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-76
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
A
B
L ENG
R ENG
F IRE FIRE E
F IIRE RE E
C
BTL 1
BTL 2
BTL 1
BTL 2
AVAIL
AVAIL
AVAIL
AVAIL
OVERHEAD
A
LEFT ENGINE FIRE PBA
B
RIGHT ENGINE FIRE PBA
FUEL MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
CTR
ON
GRAV XFR
C FUEL PANEL
CS1_CS3_2821_008
ON
Figure 38: Engine Fire PBAs and Engine Fuel Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-77
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
SYNOPTIC DISPLAY The engine fuel feed can be monitored on the FUEL synoptic page. The filter shown on the display is the main engine fuel filter and is covered in ATA 73.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-78
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
VALVE POSITION NORMAL OPERATION
9000 KG 3000 KG
Symbol
Condition
VALVE POSITION FAIL OPERATION Symbol
Open
Open
Closed
Closed
In Transition
In Transition
Invalid
Invalid
PUMPS
ENGINE FILTER Symbol
Symbol
23 °C 2000 KG
23 °C 2000 KG
Condition
Condition
Condition
Active
Engine ON - Normal
Off
Engine OFF - Normal
Failed
Engine ON or OFF - Degraded
5000 KG
Invalid Invalid Engine ON - Bypass
CS1_CS3_2821_009
APU
SYNOPTIC PAGE - FUEL
Figure 39: Fuel Synoptic Page Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-79
CS130_28_L2_EFS_ControlsIndications.fm
28 - Fuel 28-21 Engine Feed System
DETAILED DESCRIPTION LEFT ENGINE FEED The left AC boost pump can be turned on automatically or manually using the L BOOST PUMP switch on the fuel panel. The ON or OFF positions provide manual control. When the switch is in the AUTO position, the AC boost pump is controlled by the logic processed in the control and distribution cabinet (CDC). The left AC boost pump is turned on automatically when: •
APU operation on the ground
•
Fuel imbalance requiring transfer to the right wing
•
Left engine start in flight
•
Low-fuel pressure at either engine
The pump turns off automatically once the engine is running and the ejector pump is supplying pressure. The left AC boost pump is powered through the essential bus contactor located in the electronic power center (EPC) 3. If AC electrical power is lost, the ram air turbine (RAT) powers the AC ESS BUS, allowing the AC boost pump to operate. If low-fuel pressure is detected by the engine fuel pressure switch, both AC boost pumps are automatically turned on. The dual-contact switch provides logic to the CDC. It also provides status information to the data concentrator unit module cabinet (DMC). The AC boost pump pressure switch monitors the pump output and reports to the DMC. The L ENG FIRE PBA closes the engine fuel shutoff valve (FSOV) when pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-80
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
ICCP
L ENGINE FEED SHUTOFF VALVE
L Fuel SOV DC EMER BUS
Open
3
Close Ground
L ENG
FIR RE E
Engine and APU Fire Panel
L BOOST PUMP AUTO
OFF
Position
DMC 1 AND DMC 2
ON
L ENGINE FUEL PRESSURE SWITCH
CDC 1-5-17 L Fuel Pump DC ESS SSPC 3A BUS 1
To CDC 1 and CDC 2
Logic: Advanced L AC BOOST PUMP PRESSURE SWITCH
EPC 3 L Fuel Pump Fuel Panel
AC ESS BUS
10 10 10
LEGEND
EPC
CDC 1 Pump Relay Position
ARINC 429 AFDX A664 Electrical Power Center
CS1_CS3_2821_010
L AC BOOST PUMP
Figure 40: Left Engine Feed Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-81
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
RIGHT ENGINE FEED The right AC boost pump can be turned on automatically or manually using the R BOOST PUMP switch on the fuel panel. The ON or OFF positions provide manual control. When the switch is in the AUTO position, the AC boost pump is controlled by the logic processed in the CDC. The right AC boost pump is turned on automatically when: •
Fuel imbalance requiring transfer to the left wing
•
Right engine start in flight
•
Low-fuel pressure at either engine
The pump turns off automatically once the engine is running and the ejector pump is supplying pressure. The right AC boost pump is powered by AC BUS 2. If low-fuel pressure is detected by the engine fuel pressure switch, both AC boost pumps are automatically turned on. The dual-contact switch provides logic to the CDC. It also provides status information to the DMC. The AC boost pump pressure switch monitors the pump output and reports to the DMC. The R ENG FIRE PBA closes the engine FSOV when pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-82
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
ICCP
R ENGINE FEED SHUTOFF VALVE
R Fuel SOV DC EMER BUS
Open
3
Close Ground
R ENG
Position DMC 1 AND DMC 2
FIR RE E
R ENGINE FUEL PRESSURE SWITCH
Engine and APU Fire Panel
R BOOST PUMP
To CDC 1 and CDC 2
AUTO
OFF
ON
R AC BOOST PUMP PRESSURE SWITCH
Fuel Panel
Logic: Advanced LEGEND ARINC 429 AFDX A664
R AC BOOST PUMP
CS1_CS3_2821_011
CDC 2-16-1,2,3 R Fuel Pump AC BUS 2 SSPC 10A
Figure 41: Right Engine Feed Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-83
CS130_28_L3_EFS_DetailedDescription.fm
28 - Fuel 28-21 Engine Feed System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine feed system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-84
CS130_28_L2_EFS_MonitoringTests.fm
28 - Fuel 28-21 Engine Feed System
CAS MESSAGES
Table 3: STATUS Messages Table 1: CAUTION Messages
MESSAGE
MESSAGE
LOGIC
L ENG FUEL SOV Left fuel SOV failed to close/open properly. FAIL R ENG FUEL SOV Right fuel SOV failed to close/open properly. FAIL L ENG FUEL LO PRESS
Left engine feed pressure is low.
R ENG FUEL LO PRESS
Right engine feed pressure is low.
LOGIC
L BOOST PUMP ON
L boost pump is turned on via the control panel.
R BOOST PUMP ON
R boost pump is turned on via the control panel.
L BOOST PUMP OFF
L boost pump is turned off via the control panel.
R BOOST PUMP OFF
R boost pump is turned off via the control panel.
Table 4: INFO Messages Table 2: ADVISORY Messages
MESSAGE
MESSAGE
LOGIC
L BOOST PUMP FAIL
L boost pump failed to indicate high-pressure when turned on manually or automatically.
R BOOST PUMP FAIL
R boost pump failed to indicate high-pressure when turned on manually or automatically.
L FUEL EJECTOR FAIL
Left engine motive flow is faulty.
R FUEL EJECTOR Right engine motive flow is faulty. FAIL L ENG FUEL SOV CLSD
Left FUEL SOV is closed after the left engine fire switch is pressed.
R ENG FUEL SOV CLSD
Right FUEL SOV is closed after the right engine fire switch is pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
28 FUEL FAULT - Engine inlet pressure switch fault is detected. This may be ENG INLET triggered when the engine pressure switch is not high PRESS SW INOP when the fuel pump is on and providing high pressure. This may be triggered when there is a disagreement between the state of the fuel engine inlet pressure switch state as monitored by the CDCs and as monitored by the DMCs. 28 FUEL FAULT BOOST PUMP PRESS SW FAIL HI
TECHNICAL TRAINING MANUAL
For Training Purposes Only
Left and/or right boost pump pressure switch failed at high-pressure.
28-85
CS130_28_L2_EFS_MonitoringTests.fm
28 - Fuel 28-21 APU Feed System
28-21 APU FEED SYSTEM GENERAL DESCRIPTION The auxiliary power unit (APU) feed shutoff valve (SOV) isolates the APU from the fuel system. The APU feed shutoff valve opens when the APU switch is in the RUN or START position. The APU feed SOV closes when: •
Commanded OFF from the flight deck
•
APU protective shutdown occurs
•
APU fire pushbutton annunciator (PBA) is pressed
The APU is fed from the left hand engine feed line via an APU feed SOV mounted aft of the center fuel tank. The APU feed SOV is operated by a 28 VDC actuator. The APU feed SOV isolates the APU from the fuel system after APU shutdown or when commanded OFF from the flight deck. During APU ground operation, the left AC boost pump runs to provide APU feed. The arrangement of the isolation check valves between the main ejector pumps means that only the left ejector pump provides fuel to the APU while either AC pump can feed the APU. If the APU fuel burn causes an imbalance, the left and right AC pumps alternate automatically, depending on what side needs to have fuel consumed to correct imbalance.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-86
CS130_28_L2_AFS_General.fm
28 - Fuel 28-21 APU Feed System
L ENG
APU OFF
RUN
R ENG
START
PULL TURN
SOL
L DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
APU START SWITCH
SOL
R DRY BAY
M
M M
L MAIN TANK
AC
M
PS
SOL
Engine Feed Motive Flow AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
COLLECTOR TANK
COLLECTOR TANK
R SURGE TANK
M
M
SOL
APU
CS1_CS3_2821_012
LEGEND
L SURGE TANK
R MAIN TANK
Figure 42: APU Fuel Feed System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-87
CS130_28_L2_AFS_General.fm
28 - Fuel 28-21 APU Feed System
COMPONENT LOCATION The following components are installed in the APU fuel feed system: •
APU feed SOV and actuator
•
APU fuel line shroud (refer to figure 44)
APU FEED SHUTOFF VALVE AND ACTUATOR The actuator is mounted outside of the center fuel tank between RIB 0 and RIB 1R, and can be replaced without draining the fuel tank. The valve is located inside the fuel tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-88
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel 28-21 APU Feed System
APU Feed Shutoff Valve Body
APU Feed Shutoff Valve Actuator To APU
RIB 1R
RIB 0
CS1_CS3_2821_013
Rear Spar
Figure 43: APU Feed Shutoff Valve and Actuator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-89
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel 28-21 APU Feed System
APU FEED LINE, SHROUD, AND DRAIN MASTS The APU feed line runs through the fuselage from the center tank rear spar to the APU. The line is shrouded to collect any fuel that might leak from the line. Any fuel leak drains into the shroud and out one of two drain masts, located on the aft fuselage and the aft right side of the wingto-body fairing (WTBF). The center tank refuel shutoff solenoid valve also drains into the WTBF drain mast.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-90
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel
APU Fuel Drain Mast
APU Feed Line and Shroud
Refuel Solenoid Valve Drain
CS1_CS3_2821_014
28-21 APU Feed System
Figure 44: APU Fuel Line, Shroud and Drain Masts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-91
CS130_28_L2_AFS_CompLocation.fm
28 - Fuel 28-21 APU Feed System
CONTROLS AND INDICATIONS APU START SWITCH The APU feed SOV is controlled from the APU start switch located on the APU panel. APU FIRE PBA The APU FIRE PBA is located on the ENGINE and APU FIRE panel.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-92
CS130_28_L2_AFS_ControlsIndications.fm
28 - Fuel 28-21 APU Feed System
A
APU
FII R F RE E
B
OVERHEAD
BTL AVAIL
APU OFF
RUN
A ENGINE AND APU FIRE PANEL
START
CS1_CS3_2821_015
PULL TURN
B APU START SWITCH
Figure 45: APU Start Switch and Fire PBA Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-93
CS130_28_L2_AFS_ControlsIndications.fm
28 - Fuel 28-21 APU Feed System
DETAILED DESCRIPTION The APU electronic control unit (ECU) controls the APU feed SOV through the relaxed contacts of the APU feed SOV relay located in the control and distribution cabinet 1 (CDC 1). When APU switch is in the OFF position, the APU feed SOV closes. In the event of an APU fire, the APU feed SOV closes. The APU feed SOV relay energizes when the fire detection and extinguishing (FIDEX) control unit detects an APU fire or when the APU FIRE PBA is pressed. Microswitches in the APU feed shutoff valve provide valve position indication on the engine indication and crew alerting system (EICAS). The APU feed SOV can be tested through the onboard maintenance system (OMS). The APU feed SOV test is listed under ATA 49 APU tests.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-94
CS130_28_L3_AFS_DetailedDescription.fm
28 - Fuel 28-21 APU Feed System
ICCP APU Fuel SOV DC EMER BUS
FIDEX CONTROL UNIT
3 Unattended Mode
APU CDC 1 FIRE
K1 APU FIRE PBA
APU FUEL SHUTOFF VALVE Close
Close Command
APU ECU
Open
Open Command
Ground
Position
EICAS SYN DMC 1 AND DMC 2
HMU
CS1_CS3_2821_016
OMS
Figure 46: APU Feed Shutoff Valve Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-95
CS130_28_L3_AFS_DetailedDescription.fm
28 - Fuel 28-21 APU Feed System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the APU feed system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-96
CS130_28_L2_AFS_MonitoringTests.fm
28 - Fuel 28-21 APU Feed System
CAS MESSAGES Table 5: CAUTION Message MESSAGE APU FUEL SOV FAIL
LOGIC APU fuel SOV failed to close/open properly.
Table 6: ADVISORY Message MESSAGE APU FUEL SOV CLSD
LOGIC APU FUEL SOV is closed after the APU fire switch is pressed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-97
CS130_28_L2_AFS_MonitoringTests.fm
28 - Fuel 28-22 Fuel Transfer System
28-22 FUEL TRANSFER SYSTEM GENERAL DESCRIPTION The fuel transfer system has the following subsystems: •
Scavenge system
•
Automatic transfer
•
Center tank to main tank fuel transfer
•
Manual transfer -
Main tank to main tank fuel transfer
-
Main tank to center tank fuel transfer
-
Gravity transfer
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-98
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Scavenge System
Automatic Transfer FUEL TRANSFER SYSTEM
Center to Main Tank Transfer Main Tank to Main Tank
Manual Transfer
Main Tank to Center Tank
CS1_CS3_2822_010
Gravity Transfer
Figure 47: Fuel Transfer System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-99
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
SCAVENGE SYSTEM Scavenge pumps minimize the amount of unusable fuel in the main tanks. The engine-driven pump (EDP) provides motive flow to scavenge ejector pumps. The pumps are located in the sump of each main tank and continuously feed the collector tanks.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-100
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG
SOL
L DRY BAY
CENTER TANK
PS
SOL
R DRY BAY PS PS
AC
AC
PS
M
M M
L MAIN TANK
LEGEND
AC
High-Pressure Fuel Scavenge AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
COLLECTOR TANK
COLLECTOR TANK
R SURGE TANK
M
SOL
M
One-Way Check Valve PS
APU
Pressure Switch
Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
CS1_CS3_2822_005
Motor Shutoff Valve
M
SOL
L SURGE TANK
R MAIN TANK
Figure 48: Scavenge System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-101
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
AUTOMATIC TRANSFER Automatic transfer corrects fuel imbalances when the MAN XFR switch on the FUEL panel is in the OFF position. Automatic transfer is accomplished through automatic control of the AC boost pumps. If the AC boost pumps are set to AUTO and the fuel quantity computer senses an imbalance of 182 kg (400 lb) between the left and right main tanks, the AC boost pump of the heavier tank turns on automatically. Both engines receive fuel directly from the heavier tank. When the tanks are within 45 kg (100 lb) of each other, the AC boost pump is switched off. If the fuel transfer does not correct the imbalance and the imbalance reaches 364 kg (800 lb), a FUEL IMBALANCE caution message is displayed. The FUEL synoptic page displays the operating boost pump in green, and a white crossfeed arrow indicates the direction of the crossfeed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-102
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG
STATUS FUEL TOTAL FUEL FUEL USED
SOL
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
4150 KG 6000 KG
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
2 °C 2000 KG
2 °C 2150 KG
R MAIN TANK
L MAIN TANK
0 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
M
SOL
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
APU
AC
Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
M
PS
SOL
AUTO
CTR R
ON
GRAV XFR
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
ON
FUEL PANEL
CS1_CS3_2822_007
LEGEND
ON
Figure 49: Automatic Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-103
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
CENTER TANK TO MAIN TANK FUEL TRANSFER Fuel transfer from the center tank to each main wing tank, is controlled by its corresponding transfer float valve. The transfer float valve opens when the fuel level of the main tank drops below 86% of its capacity. Fuel from the center tank is automatically transferred to the main tanks by transfer ejector pumps. The pump pickup is in the center fuel tank sump. Fuel flow only occurs if the fuel quantity in the main fuel tank is low enough that the associated transfer float valve opens to allow the fuel transfer. The motive flow for the pumps is provided by the engine feed ejector pump or the AC boost pump if it is running. When the main tank level increases above 86%, the transfer float valve closes. A check valve, in the suction port of each transfer ejector prevents fuel transfer back to the center tank. This transfer operation is fully automatic and continuously replenishes the main tanks until the center tank is empty. The automatic transfer of fuel from the center tank to the wing tanks ensures that the center tank fuel is always emptied first. The center tank fuel transfer is indicated on the synoptic page by green arrows pointing outward from the center tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-104
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG
STATUS S FUEL FUE EL TOTAL FUEL FUEL USED
SOL
AIR
ELEC
FLT CTR CTRL TRLL
HYD YD
INFO FO
CB B
9000 KG 6000 KG
SOL
L DRY BAY
R DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
M
M M
L MAIN TANK
23 °C 2000 KG
R MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 2000 KG 1000 KG
R SURGE TANK
APU
SYNOPTIC PAGE - FUEL M
SOL
M
AC
Center-to-Main Tank Transfer High-Pressure Fuel Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve
Transfer Ejector
APU
Transfer Ejector
Fuel Filter Motor Shutoff Valve
M
CS1_CS3_2822_002
LEGEND
One-Way Check Valve PS
Pressure Switch
Refuel Adapter Shutoff Valve SOL
Solenoid Shutoff Valve
Figure 50: Center Tank to Main Tank Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-105
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
MANUAL FUEL TRANSFER Fuel can also be manually transferred from wing tank to opposite wing tank or to center tank through appropriate selection of fuel switches on the integrated cockpit control panel (ICCP): •
Main tank to main tank, using FUEL MAN XFER switch
•
Main tank to the center tank, using FUEL MAN XFER switch and L BOOST PUMP or R BOOST PUMP switch
•
Main to main tank by gravity, using GRAV XFR PBA
Manual fuel transfer is accomplished by connecting the engine feed line to the refuel manifold using the defuel/isolation transfer valve and the refuel shutoff valves. The defuel/isolation transfer valve and refuel shutoff valves are controlled by the fuel quantity computer (FQC). Main Tank to Main Tank Fuel Transfer Selection of the FUEL MAN XFER switch to L or R opens the defuel/isolation transfer SOV and left or right main tank refuel SOV. When the L or R BOOST PUMP switches are in AUTO, the AC boost pump in the tank the fuel is transferred from turns on. When 182 kg (400 lb) has been transferred, an EICAS FUEL MAN XFR COMPLETE advisory message is displayed. The defuel/isolation transfer SOV and the refuel SOV close and the left or right boost pump turns off.
CAUTION Do not operate the L(R) AC boost pump if the fuel quantity in the associated tank is less than 363 kg (800 lb). The boost pump can be damaged.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-106
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Right Wing Refuel Solenoid Shutoff Valve
R ENG
Defuel/Isolation Transfer Shutoff Valve
STATUS FUEL TOTAL FUEL FUEL USED
SOL
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
4100 KG 3000 KG
SOL
L DRY BAY
R DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
M
M M 23 °C 2100 KG
R MAIN TANK
L MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 2000 KG 0 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
M
SOL
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
CTR
ON
GRAV XFR
APU
LEGEND
ON
M
PS
SOL
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve
Manual Defuel Valve Lever
Center Tank Refuel Solenoid Shutoff Valve
Right Wing Refuel Shutoff Valve
Solenoid Shutoff Valve
FUEL PANEL
CS1_CS3_2822_006
AC
Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
Figure 51: Main Tank to Main Tank Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-107
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Main Tank to Center Tank Fuel Transfer When the MAN XFR switch is set to CTR, the defuel/isolation transfer SOV and the center tank refuel SOV open. The left or right boost pump operation is not automatic and must be selected ON to enable fuel transfer. A FUEL XFR CTR READY status message is displayed if the pumps are not turned on. When 364 kg (800 lb) has been transferred, an EICAS FUEL MAN XFR COMPLETE advisory message is displayed. The defuel/isolation transfer SOV and the center tank refuel SOV close. The AC boost pump switch must be returned to the AUTO or OFF position, otherwise the pump continues to run.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-108
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Left Wing Refuel Solenoid Shutoff Valve
L ENG
Right Wing Refuel Solenoid Shutoff Valve
R ENG
STATUS FUEL TOTAL FUEL FUEL USED
SOL
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
2000 KG 0 KG
SOL
L DRY BAY
R DRY BAY
CENTER TANK
PS
PS PS
AC
AC
PS
M
M M 23 °C 500 KG
R MAIN TANK
L MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 1000 KG 500 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
M
APU
SOL
L
Center Tank Refuel Shutoff Valve
LEGEND
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
CTR
ON
GRAV XFR ON
M
PS
SOL
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
Manual Defuel Center Tank Refuel Solenoid Shutoff Valve Valve Lever
FUEL PANEL
NOTE Aircraft shown on ground.
CS1_CS3_2822_009
AC
Engine Feed AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
R
L BOOST PUMP
Figure 52: Main Tank to Center Tank Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-109
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
Gravity Transfer The GRAV XFR PBA transfers fuel between the main tanks to correct the imbalances between the wing tanks. When the GRAV XFR PBA is pressed, it opens the gravity crossflow SOV. The gravity crossflow SOV is normally closed to provide isolation between both wing tanks. The GRAV XFR PBA is hardwired to the gravity crossflow shutoff valve. When the SOV is open, the fuel level in the main tanks is allowed to equalize through gravity. The gravity crossflow is displayed on the FUEL synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-110
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
L ENG
R ENG TOTAL FUEL FUEL USED
SOL
AIR
ELEC
FLT CTRL
HYD
INFO
CB
4000 KG 0 KG
SOL
L DRY BAY PS
R DRY BAY
CENTER TANK
PS PS
AC
AC
PS
M
M M
L MAIN TANK
23 °C 2000 KG
R MAIN TANK
COLLECTOR TANK
L SURGE TANK
COLLECTOR TANK
23 °C 2000 KG 0 KG
APU
R SURGE TANK
SYNOPTIC PAGE - FUEL FUEL
M
MAN XFR OFF
L
SOL
APU
AC
Gravity Transfer AC Boost Pump Ejector Pump Flapper Check Valve Float Valve Fuel Filter
M
PS
SOL
Motor Shutoff Valve One-Way Check Valve Pressure Switch Refuel Adapter Shutoff Valve Solenoid Shutoff Valve
R BOOST PUMP
AUTO
OFF
LEGEND
R
L BOOST PUMP ON
AUTO
CTR R
OFF
ON
GRAV XFR
Gravity Crossflow Shutoff Valve
Gravity Feed Manifold
ON
FUEL PANEL
CS1_CS3_2822_004
M
Figure 53: Gravity Fuel Transfer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-111
CS130_28_L2_FTS_General.fm
28 - Fuel 28-22 Fuel Transfer System
COMPONENT LOCATION The scavenge system consists of the following component: •
Scavenge ejector pump
The center tank to main tank transfer system consists of the following components: •
Transfer float valve
•
Transfer ejector pump
The gravity fuel transfer system consists of the following component: •
Gravity crossflow shutoff valve
SCAVENGE EJECTOR PUMP The scavenge ejector pumps are mounted on RIB 7 of each wing tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-112
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
RIB 5L Collector Tank Motive Flow Line
NOTE SCAVENGE EJECTOR PUMP
Left wing shown, right wing similar.
CS1_CS3_2821_004
RIB 7L
Figure 54: Scavenge Ejector Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-113
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
TRANSFER FLOAT VALVE This valve is attached to the center tank transfer discharge line at RIB 16. TRANSFER EJECTOR PUMP Each of the two transfer systems contains an ejector pump installed at RIB 4 of the center fuel tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-114
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
RIB 16L
Transfer Float Valve
RIB 4L
Transfer Ejector Pump
TRANSFER FLOAT VALVE
NOTE Left wing shown. Right wing similar.
TRANSFER EJECTOR PUMP
CS1_CS3_2822_015
Transfer Ejector Pump Pickup
Figure 55: Center Tank to Main Tank Fuel Transfer Ejector Pump and Float Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-115
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
GRAVITY CROSSFLOW SHUTOFF VALVE The gravity crossflow shutoff valve is located inboard of RIB 2R on the rear spar.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-116
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
Gravity Crossflow Shutoff Valve
RIB 2
Gravity Crossflow Shutoff Valve Actuator
Gravity Valve Pipe RIB 1R
CS1_CS3_2822_003
Rear Spar
Figure 56: Gravity Crossflow Shutoff Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-117
CS130_28_L2_FTS_CompLocation.fm
28 - Fuel 28-22 Fuel Transfer System
CONTROLS AND INDICATIONS Fuel transfer is controlled from the FUEL panel on the right inboard integrated cockpit control panel (ICCP). The MAN XFR switch is used to manually transfer fuel between the tanks. The MAN XFR switch is a four-position rotary switch located on the FUEL panel. The normal position of the switch is OFF. In the OFF position, automatic imbalance correction is enabled. The GRAV XFR PBA opens the gravity crossflow SOV. The gravity crossflow SOV is hardwired to the valve.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-118
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
FUEL MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
ON
AUTO
R CTR
ON
GRAV XFR ON
CS1_CS3_2823_010
FUEL PANEL
Figure 57: Fuel Transfer Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-119
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
SYNOPTIC PAGE Fuel transfer operations are shown on the FUEL synoptic page and the EICAS page. The arrows show the direction of transfer during an automatic transfer. The refuel valves are only shown in amber if they fail during a manual transfer operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-120
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
9500 KG 3000 KG
VALVE POSITION FAIL OPERATION
VALVE POSITION NORMAL OPERATION Symbol
Symbol
Condition
Condition
Symbol
Condition
Symbol
Condition
Open
Open
Active
Active
Closed
Closed
Off
Failed/Low
In Transition
In Transition
Failed
No Transfer
Invalid
Invalid
Invalid
Invalid
FLOW LINES Symbol
23 °C 2500 KG
CENTER TO WING XFERS
PUMPS
TRANSFER FUNCTION ACTIVE
Condition
Symbol
Fuel Flow
23 °C 2000 KG
Condition To R Wing
No Fuel Flow
To L Wing
Invalid Fuel Flow
5000 KG
TOTAL FUEL (KG)
APU
2500
5000 EICAS PAGE
Refuel Shutoff Valves SYNOPTIC PAGE - FUEL
9500 2000 CS1_CS3_2822_008
TOTAL FUEL FUEL USED
Figure 58: Fuel Synoptic Page Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-121
CS130_28_L2_FTS_ControlsIndications.fm
28 - Fuel 28-22 Fuel Transfer System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the fuel transfer system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-122
CS130_28_L2_FTS_MonitoringTests.fm
28 - Fuel 28-22 Fuel Transfer System
CAS MESSAGES Table 9: STATUS Messages
Table 7: CAUTION Messages MESSAGE
LOGIC
MESSAGE
LOGIC
FUEL IMBALANCE
Fuel imbalance between left and right tank greater than 364 kg (800 lb).
FUEL MAN XFR TO L
Manual XFR TO L WING is selected on the fuel control panel.
FUEL CTR XFR FAIL
Fuel transfer out of center tank has failed, center tank fuel may become unusable.
FUEL MAN XFR TO R
Manual XFER TO R WING is selected on the fuel control panel.
FUEL MAN XFR FAIL
Manual transfer failed to perform the commanded operation or to stop the transfer.
FUEL MAN XFR TO CTR
MAN XFR to the center tank is selected and either L or R boost pump has been turned on.
FUEL GRAV XFR GRAVITY SOV has been commanded open via the ON overhead panel.
Table 8: ADVISORY Messages MESSAGE FUEL MAN XFR COMPLETE
LOGIC Wing-to-wing target fuel transfer of 182 kg (400 lb) or wing-to-center target fuel transfer of 364 kg (800 lb) has completed.
FUEL XFR CTR READY
MAN XFR to the center tank is selected but no pump has been selected.
FUEL GRAV XFR FUEL GRAV SOV failed to open or close. FAIL FUEL CTR XFR FAULT
Fuel transfer from center to the left or right tank has failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28-123
CS130_28_L2_FTS_MonitoringTests.fm
28 - Fuel 28-40 Fuel Indicating System
28-40 FUEL INDICATING SYSTEM GENERAL DESCRIPTION The fuel indicating system includes a fuel quantity computer (FQC), remote data concentrators (RDC), fuel probes, and temperature sensors. The purpose of the fuel probes is to measure the quantity of fuel in the tank. The probes are located strategically throughout the fuel tank to provide the total quantity of fuel in the tank. Compensator probes provide fuel dielectric value used to compensate the fuel quantity for changes in fuel temperature. Low-level fuel probes provide a warning when approximately 30 minutes fuel remain. The high-level fuel probes prevent an overfill of the fuel tanks. Temperature sensors in each main tank provide the fuel temperature input to the fuel indicating system. The main fuel tank temperatures are displayed on the FUEL synoptic page. The fuel probes and compensators in each tank are organized in two arrays called lane 1 and 2. Each array is connected to a processor in the RDC. Each RDC has two independent and identical processors that convert the fuel probe data into digital format. The RDCs send the tank data digitally to the FQC. The FQC has dual independent and identical processors for processing the data received from the RDC units. The FQC performs all of the computations associated with the fuel system. The FQC provides fuel quantity and temperature information to the engine indication and crew alerting system (EICAS). The FQC and RDC software is field-loadable.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-124
CS130_28_L2_FIS_General.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL FUEL
MAN MANXFR XFR OFF OFF
LL
REFUEL / DEFUEL PANEL
BOOSTPUMP PUMP LL BOOST OFF
OFF
AUTO AUTO
LEFT
ON
RR
CENTER
RIGHT
PRESEL INCR
BOOSTPUMP PUMP RR BOOST
ON
R CTR CTR
ON
GRAV XFR GRAV XFR
OFF
AUTO AUTO
OFF
OFF
ON
TOTAL
PRESEL
DECR
ON LEFT
CENTER
RIGHT OPEN
ON ON
DEFUEL
AUTO
START
CLOSE
MANUAL
FUEL PANEL
REFUEL
MANUAL STOP/SOV TEST
REFUEL/DEFUEL CONTROL PANEL Fuel Door Switch
FQC Primary
REFUEL SHUTOFF VALVES
Backup
DC EMER BUS DC ESS BUS 1
Fuel Quantity Computer Channel 1
EICAS
28 VDC
Fuel Quantity Computer Channel 2
DC ESS BUS 2
SYN OMS HMU
CENTER TANK
Fuel Quantity Probes Array 2
Fuel Quantity Probes Array 1
Right RDC
RIGHT MAIN TANK
RIGHT MAIN TANK TEMPERATURE SENSOR
CS1_CS3_2841_009
LEFT MAIN TANK
Fuel Quantity Probes Array 2
Remote Data Concentrator ARINC 429 CAN BUS CAN BUS RS-422
Center RDC Fuel Quantity Probes Array 1
RDC
Fuel Quantity Probes Array 2
LEGEND
Left RDC Fuel Quantity Probes Array 1
LEFT MAIN TANK TEMPERATURE SENSOR
Figure 59: Fuel Quantity Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-125
CS130_28_L2_FIS_General.fm
28 - Fuel 28-40 Fuel Indicating System
COMPONENT LOCATION The fuel quantity indication system components consist of: •
Fuel probes
•
Temperature sensor
•
Remote data concentrator (refer to figure 61)
•
Fuel quantity computer (refer to figure 62)
FUEL PROBES There are 58 fuel probes installed in the aircraft fuel tanks. TEMPERATURE SENSORS There is one temperature sensor installed in each of the aircraft main fuel tanks. The temperature sensor is mounted in a housing on RIB 7.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-126
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
LEGEND
CS1_CS3_2822_011
Temperature Sensor Lane 1 Fuel Probe Lane 2 Fuel Probe Lane 2 Fuel Probe/Compensator Lane 1 Fuel Probe/Compensator Lane 1 Fuel Probe (Derived Low Level) Lane 2 Fuel Probe (Derived Low Level) Lane 1 Fuel Probe (Derived High Level) Lane 2 Fuel Probe (Derived High Level)
Figure 60: Fuel Quantity Probes and Temperature Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-127
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
REMOTE DATA CONCENTRATOR The wing tank remote data concentrators (RDCs) are located on the front spar outboard of the engines. The center RDC is mounted in the wing-to-body fairing (WTBF) on FRAME 49.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-128
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
Center RDC
Frame 49
REMOTE DATA CONCENTRATOR (LEFT WING SHOWN, RIGHT WING SIMILAR)
CENTER REMOTE DATA CONCENTRATOR (LEFT WHEEL BIN REMOVED)
CS1_CS3_2841_003
Left Wing RDC
Figure 61: Remote Data Concentrator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-129
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL QUANTITY COMPUTER The fuel quantity computer (FQC) is located in the mid equipment bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28-130
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
CS1_CS3_2841_004
MID EQUIPMENT BAY AFT RACK
FUEL QUANTITY COMPUTER
Figure 62: Fuel Quantity Computer Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-131
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL PROBES The tank units are two element concentric metal tube assemblies open at each end to allow fuel in. The probes are mounted on brackets attached to ribs and located throughout the tank. The length of each tank unit is determined by the height of the fuel tank at its particular location. In addition to providing quantity information, some fuel quantity probes perform other functions, as follows: •
High-level probes - There are two capacitance type fuel high-level probes installed in each main fuel tank and four in the center tank. The high-level probes are mounted outboard of RIB 20 in the main tanks, and at RIB 6 on the left and right sides of the center tank
•
Low-level probes - There are four capacitance type low-level fuel probes in each wing tank. Two fuel probes are located in each main tank outboard of RIB 7, and two are in each collector tank on the inboard side of RIB 7
•
Compensator probes - The compensator is a multiple concentric metal tube assembly mounted near the bottom of each main tank. In this position, the compensator is always covered with fuel until each tank almost empty. Any change in capacitance is the result of change of dielectric constant of the fuel. The compensator probes are located at RIB 5 of the main tanks and RIB 0 of the center tank
TEMPERATURE SENSOR The temperature sensor attaches to the housing with a bayonet fitting. The sensor is accessed from the dry bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-132
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
FUEL PROBE COMPENSATOR LOW-LEVEL FUEL PROBE
FUEL PROBE
TEMPERATURE SENSOR
CS1_CS3_2822_013
HIGH-LEVEL FUEL PROBE
Figure 63: Fuel Probes and Temperature Sensor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-133
CS130_28_L2_FIS_CompLocation.fm
28 - Fuel 28-40 Fuel Indicating System
DETAILED COMPONENT INFORMATION FUEL QUANTITY COMPUTER INTERFACE The fuel quantity computer interfaces with other aircraft systems through the data concentrator unit module cabinet (DMC). The FQC provides fuel quantity and fuel temperature information to the DMC via ARINC 429 BUSES. The FQC provides: •
Defuel and refuel valve control
•
Fuel quantity and temperature related CAS messages
•
High and low fuel level warning
•
Fuel imbalance warning
•
Automatic wing fuel transfer
•
Communication with the DMC
•
Onboard maintenance system communication
•
Communication with the RDCs
Most fuel system electrical components interface directly with the electrical system. The control signals are derived from the direct discrete signals, such as the fire handle, or indirect signals via an ARINC 429 BUS. The operational commands from these systems control the operation of the fuel system pumps and valves. The operating status of the pumps and valves are fed to the DMC and the electrical system. The fuel system components interface with the DMC to provide information to the engine indication and crew alerting system (EICAS), FUEL synoptic page display, and the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-134
CS130_28_L3_FIS_DetailedCompInfo.fm
28 - Fuel 28-40 Fuel Indicating System
EPC
FUEL
Gravity Crossflow Valve
MAN XFR OFF
L
R
L BOOST PUMP
R BOOST PUMP
OFF
OFF
AUTO
AUTO
ON
R CTR
L ENG Fuel Press
L ENG LO Pressure
R ENG Fuel Press
L ENG LO Pressure
Position
ON
Pressure
GRAV XFR ON
FUEL PANEL
APU Fire PBA
L AC Boost Pump
R AC Boost Pump Pressure
FIDEX Control Unit APU Switch On APU Emergency Stop
APU Electronic Control Unit
Relay K1 CDC 1
APU Feed Shutoff Valve DMC 1 and DMC 2
Position
L ENG Run Switch
L EEC
R ENG Run Switch
R EEC
SYN
L ENG Fire PBA
L ENG FEED SOV
OMS
EICAS
Position HMU
R ENG FEED SOV
LEGEND
REFUEL / DEFUEL PANEL LEFT
ON
Position
OFF
TOTAL LEFT
CDC EPC RDC SOV
Control and Distribution Cabinet Electrical Power Center Remote Data Concentrator Shutoff Valve A664 ARINC 429 CAN A CAN B RS-422
L Refuel SOV
RIGHT
DEFUEL
AUTO
Defuel/ Isolation Transfer SOV
PRESEL INCR
DECR
START
CLOSE
R RDC
C Refuel SOV
CENTER
RIGHT
PRESEL
OPEN
C RDC L RDC
CENTER
MANUAL
R Refuel SOV
REFUEL
STOP/SOV TEST
FUEL QUANTITY COMPUTER
Position CS1_CS3_2841_006
R ENG Fire PBA
Figure 64: Fuel Quantity Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-135
CS130_28_L3_FIS_DetailedCompInfo.fm
28 - Fuel 28-40 Fuel Indicating System
CONTROLS AND INDICATIONS The fuel quantity indication is displayed on the FUEL synoptic page. The fuel quantity, fuel used, and temperature information are also displayed on the EICAS page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-136
CS130_28_L2_FIS_ControlsIndications.fm
28 - Fuel 28-40 Fuel Indicating System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
9000 KG 0 KG
QUANTITY Symbol
23 °C 2000 KG
Condition
Symbol
3000 KG
Normal
1600 KG
Low/Overfilled/ Imbalance/Trapped
è½ KG
23 °C 2000 KG
TEMPERATURE
22 °C è44 °C è½ °C
Invalid
TOTAL FUEL (KG) 2000
5000 KG
5000
Condition Normal Low or High Invalid
9000 2000
EICAS PAGE
CS1_CS3_2841_002
APU
SYNOPTIC PAGE - FUEL
Figure 65: Fuel Controls and Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-137
CS130_28_L2_FIS_ControlsIndications.fm
28 - Fuel 28-40 Fuel Indicating System
DETAILED DESCRIPTION The fuel quantity system consists of arrays of AC capacitance probe units, located in the left wing, right wing and center fuel tanks. Each fuel tank uses an RDC to isolate the tank sensors from the dual-channel FQC. The RDCs provide information digitally to the FQC via two CAN BUSES. The FQC interfaces with a refuel/defuel control panel using an RS-422 BUS. Each RDC has two lanes for redundancy. The RDC lanes 1 and 2 have independent and identical processors that receive analog data from the fuel probes and sensors. An interprocessor connects the lanes within the RDC. This ensures that all the data from all the tank probes is available in either lane. Each lane is broken down into groups of fuel probes. An excitation generator in the RDC provides a separate excitation signal to each group. By grouping the fuel probes, the total energy in each circuit is very low. This reduces ignition source hazards. One fuel probe compensator is located in each tank. Temperature sensors provide the fuel temperature of the left and right wing tank fuel. To ensure availability of data to FQC from either wing compensator probe or temperature sensor in case of a CAN BUS failure, the temperature sensor in left tank is connected to RDC lane 2 while the temperature sensor in the right tank is connected to RDC lane 1. The compensator in the left tank is connected to RDC lane 2 and the compensator in the right tank is connected to RDC lane 1. The FQC has two identical and independent channels for processing the fuel quantity data. Each channel receives data digitally from the RDCs. The FQC uses the fuel quantity information to control the refuel/defuel operations and automatic fuel balancing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The FQC channels communicate with each other over an internal bus. The internal bus allows sharing of any input data from the other channel in case of input failures. Each FQC channel sends data to the L and R DMCs over ARINC 429 BUSES. In normal operation, the DMCs use the data from FQC channel 1. Channel 2 data is used when channel 1 fails. The FQC continuously monitors the fuel quantities in each tank and provides annunciation of potentially hazardous conditions. The low-level fuel probes in each wing tank are located at a level to provide an output at 30 minutes fuel remaining, which is approximately 443.2 kg (975 lb). The FQC provides two independent low-level indications to the DMC over ARINC 429 BUSES and discrete lines. In the event of an ARINC 429 BUS failure, the discrete signals are used for display on the EICAS for L or R FUEL LO QTY messages. The high-level fuel probes provide a signal when the fuel level exceeds 98% of the tank volume. During refueling, a high fuel level condition shuts down the refuel system and annunciates a TANK HIGH LEVEL on the refuel/defuel control panel (RDCP). The flight crew has no indication of a leak other than information derived from fuel quantity data. A FUEL LEAK SUSPECT caution message provides indication of a possible fuel leak if any of the following conditions are present: •
The fuel system suspects a wing tank overfill condition is present
•
The electronic engine control detects excessive fuel consumption by comparing both engines
•
The FMS calculated fuel used and indicated fuel onboard is different
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-138
CS130_28_L3_FIS_DetailedDescription.fm
28 - Fuel 28-40 Fuel Indicating System
FMS
DMC 1 AND DMC 2
EICAS
REFUEL/DEFUEL CONTROL PANEL
28 VDC
DC EMER BUS Fuel Quantity Computer Channel 1
FUEL QUANTITY COMPUTER
SYN
Fuel R Low Level
Fuel L Low Level
LEFT EEC AND RIGHT EEC
Fuel Quantity Computer Channel 2
DC ESS BUS 1
OMS HMU
DC ESS BUS 2
L REFUEL VALVE
CENTER TANK RDC
LEFT WING RDC
Lane 2
Lane 1
Lane 2
Lane 1
LEFT WING LANE 1 CENTER TANK LANE 1 AND COMPENSATOR
RIGHT WING LANE 1 AND COMPENSATOR
CENTER TANK LANE 2
LEGEND Remote Data Concentrator ARINC 429 CAN BUS CAN BUS Discrete Excitation RS-422
R REFUEL VALVE
RIGHT WING LANE 2
LEFT WING LANE 2 AND COMPENSATOR
RDC
Lane 2
C REFUEL VALVE
LEFT MAIN TANK TEMPERATURE SENSOR
RIGHT MAIN TANK TEMPERATURE SENSOR
CS1_CS3_2841_005
Lane 1
RIGHT WING RDC
DEFUEL/ ISOLATION TRANSFER VALVE
Figure 66: Fuel Quantity System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
28-139
CS130_28_L3_FIS_DetailedDescription.fm
28 - Fuel 28-40 Fuel Indicating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the fuel quantity indication system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-140
CS130_28_L2_FIS_MonitoringTests.fm
28 - Fuel 28-40 Fuel Indicating System
CAS MESSAGES
Table 12: INFO Messages Table 10: CAUTION Messages
MESSAGE
MESSAGE
28 FUEL FAULT - Fuel computer channel fault is detected. COMPUTER REDUND LOSS
LOGIC
L FUEL LO QTY
Low-level detected in left wing tank.
R FUEL LO QTY
Low-level detected in right wing tank.
FUEL TANK LO TEMP
Fuel temp is near Jet A fuel freezing point (-37 ºC).
FUEL TANK HI TEMP
Fuel temp above the operating limit.
FUEL COLLECTOR LO LVL
Fuel in the collector is low - suspect the transfer ejector is inoperative for the left or right tank.
FUEL LEAK SUSPECT
This may be caused by fuel leaking overboard due to: • A wing tank fuel overfill condition from the center transfer float valve stuck open • A manual transfer refuel solenoid valve failed to close Or this message may be displayed when: • The EEC suspects a leak at the engine based upon comparison between the L and R engine fuel consumptions • The FMS detects a suspected fuel leak triggered by a difference in fuel onboard compared to the FMS calculated fuel consumption while the engines are running
Table 11: ADVISORY Messages MESSAGE
LOGIC
FUEL Fuel quantity computer has failed or loss of both COMPUTER FAIL ARINC 429 channels from FQC to DMC. FUEL FAULT
Loss of redundant or non-critical function for the fuel system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
28 FUEL FAULT - Loss of fuel quantity in the left tank. L WING RDC INOP 28 FUEL FAULT L WING RDC REDUND LOSS
Fuel left wing remote data concentrator fault is detected.
28 FUEL FAULT R WING RDC INOP
Loss of fuel quantity in the right tank.
28 FUEL FAULT R WING RDC REDUND LOSS
Fuel right wing remote data concentrator fault is detected.
28 FUEL FAULT CTR WING RDC INOP
Loss of fuel quantity in the center tank.
28 FUEL FAULT CTR WING RDC REDUND LOSS
Fuel center tank remote data concentrator fault is detected.
28 FUEL FAULT FUEL GAUGING SNSR INOP
Fuel gauging sensor (probe/compensator) fault is detected.
28 FUEL FAULT FUELING DOOR OPEN
Fuel door is not CLSD prior to flight.
28 FUEL FAULT FUEL KG-LB MISCOMPARE
Fuel system refuel panel and cockpit primary page use different units (kg/lb).
TECHNICAL TRAINING MANUAL
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28 - Fuel 28-40 Fuel Indicating System
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-40 Fuel Indicating System
Table 12: INFO Messages MESSAGE
LOGIC
28 FUEL FAULT DEFUEL/XFR SOV INOP
DEFUEL/XFER SOV has failed when the actual reported SOV position is different from the commanded position, including conflicting reported position.
28 FUEL FAULT CONFIG STRAPPING INOP
Fuel gauging computer or fuel remote data concentrator has strapping fault(s).
28 FUEL FAULT SOFTWARE LOAD INOP
Fuel gauging computer software load fault is detected.
28 FUEL FAULT FUEL TEMP SNSR INOP
One fuel temperature sensor has failed.
28 FUEL FAULT GAUGING SNSR SHORT CIRCUIT
Gauging sensor short circuit of fuel probe or wiring.
28 FUEL FAULTBOOST PUMP PRESS SW FAIL HI
Left and/or right boost pump pressure switch is failed to high-pressure.
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TECHNICAL TRAINING MANUAL
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CS130_28_L2_FIS_MonitoringTests.fm
28 - Fuel 28-40 Fuel Indicating System
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM Each channel of the FQC is tested using the OMS. The FQC enters the OMS test mode when the aircraft is weight-on-wheels (WOW) and the parking brake is set. The OMS provides a system parameter page that displays the real time reported capacitance and temperature sensor data as well as status and other fuel system related faults from the ARINC 429 data. Low-Level Discrete Output Test When the OMS commands a low-level discrete output test, the FQC channel sets the FUEL L LOW LEVEL and then the FUEL R LOW LEVEL discrete outputs true for 3 seconds. The FQC also compares the DMC reported feedback of the FUEL L LOW LEVEL and FUEL R LOW LEVEL discrete outputs to the FQC commanded state. This test checks that the low-level discrete outputs operate properly, and verifies proper wiring and communication with the DMCs. The FQC reports the pass/fail status of this test to the OMS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28-144
CS130_28_L2_FIS_PracticalAspects.fm
28 - Fuel 28-40 Fuel Indicating System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-LOW LEVEL DISCRETE OUTPUT TEST
001 /004 FUEL LEVEL OUTPUT TEST
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2841_001
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 67: Low-Level Discrete Output Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28-145
CS130_28_L2_FIS_PracticalAspects.fm
28 - Fuel 28-40 Fuel Indicating System
Additional Functions The fuel quantity computer provides the ability to clear non-volatile memory (NVM) faults and report the configuration from the OMS test screen selection. These initiated functions include: •
The CLEAR NVM function clears faults stored in the FQC NVM
•
The REPORT CONFIGURATION function reports the configuration data of the fuel quantity system. The FQC, RDCs, and RDCP report their configuration data to the OMS
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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28 - Fuel 28-40 Fuel Indicating System
MAINT MENU
RETURN TO LRU/SYS OPS
FQC CHA-CLEAR NVM
001 /004 CLEAR NVN TEST
* NOTE * KEEP A/C FUEL PUMPS OFF
PRECONDITIONS
1 2 3 4
-
MAKE MAKE MAKE MAKE
SURE SURE SURE SURE
AIRCRAFT IS WOW AIRCRAFT PARK BRAKE IS SET RDCP ACCESS DOOR IS CLOSED NO REFUEL/DEFUEL OPERATIONS ACTIVE TEST INHIBIT/INTERLOCK CONDITIONS
TEST ENABLED / INHIBITED STATUS: YYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
PRESS 'CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_2823_020
PRESS 'CONTINUE' BUTTON TO GO TO THE NEXT PAGE
Figure 68: Clear NVM Function Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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28 - Fuel 28-40 Fuel Indicating System
Page Intentionally Left Blank
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ATA 30 - Ice and Rain Protection
BD-500-1A10 BD-500-1A11
30 - Ice and Rain Protection
Table of Contents 30-81 Ice Detection System....................................................30-2 General Description .........................................................30-2 Component Location ........................................................30-4 Ice Detector..................................................................30-4 Detailed Description ........................................................30-6 Monitoring and Tests ........................................................30-8 CAS Messages ............................................................30-9 Initiated Built-In Test........................................................30-10 30-12 Wing Anti-Ice System .................................................30-12 General Description ........................................................30-12 Wing Anti-Ice Ducting ................................................30-14 Component Location .....................................................30-16 Wing Anti-Ice Valves..................................................30-16 Venturis......................................................................30-16 Telescopic Ducts........................................................30-16 Piccolo Tubes ............................................................30-16 Wing Anti-Ice Temperature Sensors..........................30-16 Pressure Sensors ......................................................30-16 Integrated Air System Controllers ..............................30-18 Controls and Indications .................................................30-20 Anti-Ice Control Panel ................................................30-20 Wing Anti-Ice indications............................................30-22 Detailed Description .......................................................30-24 Integrated Air System Controllers ..............................30-24 Wing Anti-Ice Operation.............................................30-26 Wing Anti-Ice Activation Control ................................30-28 Monitoring and Tests .....................................................30-30 System Test ...............................................................30-30 CAS Messages ..........................................................30-32 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Practical Aspects ........................................................... 30-36 Wing Anti-Ice Valve................................................... 30-36 30-22 Cowl Anti-Ice System................................................. 30-38 General Description ...................................................... 30-38 Component Location ..................................................... 30-40 Cowl Anti-Ice Valves ................................................. 30-40 Pressure Sensor ....................................................... 30-40 Fan Cowl Temperature Sensor ................................. 30-40 Swirl Nozzle .............................................................. 30-40 Detailed Component Information ................................... 30-42 Swirl Nozzle .............................................................. 30-42 Controls and Indications ................................................ 30-44 Indications ................................................................. 30-46 Detailed Description ...................................................... 30-48 Temperature and Pressure Sensing ......................... 30-48 Monitoring and Tests ..................................................... 30-50 CAS Messages ......................................................... 30-51 Practical Aspects ............................................................ 30-52 Cowl Anti-Ice Valves ................................................. 30-52 34-10 Probe Heat System.................................................... 30-54 General Description ...................................................... 30-54 Component Location ..................................................... 30-56 Angle-of-Attack Vane ................................................ 30-56 Air Data System Probe.............................................. 30-56 Total Air Temperature Probe..................................... 30-56 Static Inverter ............................................................ 30-56 Controls and Indications ................................................ 30-58 Detailed Description ..................................................... 30-60 Air Data System Probe.............................................. 30-60
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-i
CS130_M5_30_IceRainTOC.fm
30 - Ice and Rain Protection
Monitoring and Tests .....................................................30-62 CAS Messages ..........................................................30-63 30-41 Windshield and Side Window Heating System...........30-68 General Description .......................................................30-68 Component Location .....................................................30-70 Windshield .................................................................30-70 Side Window ..............................................................30-70 Windshield Ice Protection Controllers ........................30-70 Controls and Indications ................................................30-72 Detailed Description .......................................................30-74 Monitoring and Tests .....................................................30-76 CAS Messages ..........................................................30-77 30-42 Windshield Wipers System.........................................30-78 General Description .......................................................30-78 Component Location .....................................................30-80 Wiper Motor ...............................................................30-80 Wiper Arm ..................................................................30-80 Wiper Blade ...............................................................30-80 Controls and Indications .................................................30-82
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-ii
CS130_M5_30_IceRainTOC.fm
30 - Ice and Rain Protection
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7:
Ice Detection System ............................................30-3 Ice Detector ...........................................................30-5 Ice Detection System Schematic...........................30-7 Ice Detection System Test ..................................30-11 Wing Anti-Ice System ..........................................30-13 Wing Anti-Ice Air Supply......................................30-15 Wing Anti-Ice System Leading Edge Components.................................30-17 Figure 8: Integrated Air System Controllers .......................30-19 Figure 9: Wing Anti-Ice Controls.........................................30-21 Figure 10: Wing Anti-Ice Indications.....................................30-23 Figure 11: Integrated Air System Controller .........................30-25 Figure 12: Wing Anti-Ice System Schematic ........................30-27 Figure 13: Wing Anti-Ice Activation Advanced Logic............30-29 Figure 14: Wing Anti-Ice System Test ..................................30-31 Figure 15: Wing Anti-Ice Valve .............................................30-37 Figure 16: Cowl Anti-Ice System ..........................................30-39 Figure 17: Cowl Anti-Ice System Component Location ........30-41 Figure 18: Swirl Nozzle.........................................................30-43 Figure 19: Cowl Anti-Ice System Controls............................30-45 Figure 20: Cowl Anti-Ice System Indications ........................30-47 Figure 21: Cowl Anti-Ice System Detailed Description .........30-49 Figure 22: Cowl Anti-Ice Valves ...........................................30-53 Figure 23: Probe Heat System .............................................30-55 Figure 24: Probe Heat System Components........................30-57 Figure 25: Probe Heat Control..............................................30-59
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 26: Probe Heat Electrical Interface Detailed Description............................................ 30-61 Figure 27: Windshield and Window Heat System................ 30-69 Figure 28: Windshield, Side Window, and Windshield Ice Protection Controllers................. 30-71 Figure 29: Window Heat Panel ............................................ 30-73 Figure 30: Windshield and Window Heat System Schematic ..................................... 30-75 Figure 31: Windshield Wiper System................................... 30-79 Figure 32: Windshield Wiper Motor, Wiper Arm, and Wiper Blade ................................................. 30-81 Figure 33: Windshield Wipers System Controls................... 30-83
TECHNICAL TRAINING MANUAL
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30-iii
CS130_M5_30_IceRainLOF.fm
ICE AND RAIN PROTECTION - CHAPTER BREAKDOWN ICE AND RAIN PROTECTION CHAPTER BREAKDOWN
Windshield and Window Heat System
Ice Detection System
1
5
Windshield Wiper System
Wing Anti-Ice System
2
Cowl Anti-Ice System
3
Probe Heat System
4
6
30 - Ice and Rain Protection 30-81 Ice Detection System
30-81 ICE DETECTION SYSTEM CAUTION GENERAL DESCRIPTION The detection system detects the supercooled liquid water conditions present in the clouds that can cause ice formation on the aircraft.
As a safety precaution do not touch the probes with bare hands, as they could be hot and inflict burns.
The ice detectors provide independent measurements of icing conditions. The system is comprised of two magnetostrictive ice detectors. The ice detector is a one-piece, vibrating sensing element (probe) type device. When ice accumulates on the probe, the drop in frequency triggers the ice detector to generate discrete signals (ice). In addition, heaters inside the probe and strut automatically come on to melt the ice, then turn off. With the heaters off, ice is again allowed to build up. The cycle of accumulation and melting is repeated until ice is no longer present. The discrete signals are switched off 2 minutes later. The detectors operate using 115 VAC electrical power and are always powered when the aircraft is powered. The left ice detector is powered by the AC BUS 1 and the right is powered by the AC BUS 2. When ice is detected by either detector, discrete signals are automatically sent via the data concentrator unit module cabinet (DMC) to the wing anti-ice (WAIS) system and the cowl anti-ice system (CAIS) to provide anti-icing. It also signals the three primary flight control computers (PFCCs) to modify the stall warning protection. As a backup the ice detection signal is also sent to the WAIS through the control and distribution cabinets (CDCs). The ice signals and ice detector status information are sent to the engine indication and crew alerting system (EICAS) and the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-2
CS130_30_L2_IDS_General.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
LEFT ICE DETECTOR
RIGHT ICE DETECTOR
LEFT and RIGHT CAIS
Test
AC BUS 1
Test
AC BUS 2
EICAS Ice Detector OK
Ice Detector OK DMC 1
OMS
Ice Detected 1 and Ice Detected 2
DMC 2 Ice Detected 1 and Ice Detected 2
PFCCs
WAIS
Ice Detected 3
CDCs
Ice Detected 3
LEGEND Cowl Anti-Ice System Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Primary Flight Control Computer Wing Anti-Ice System ARINC 429 Discrete
CS1_CS3_3081_013
CAIS CDC DMC PFCC WAIS
Figure 1: Ice Detection System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-3
CS130_30_L2_IDS_General.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
COMPONENT LOCATION There are two identical ice detectors installed in the system. ICE DETECTOR The ice detectors are located on each side of the forward fuselage, one is on the left, and the other on the right.
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30-4
CS130_30_L2_IDS_ComponentLocation.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
Mounting Flange
Strut
Probe
Air Flow Direction
NOTE Left ice detector shown. Right ice detector similar.
CS1_CS3_3080_002
ICE DETECTOR
Figure 2: Ice Detector Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-5
CS130_30_L2_IDS_ComponentLocation.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
DETAILED DESCRIPTION The ice detector is a one-piece, vibrating sensing element (probe) type device. An internal electrical oscillator circuit drives the sensing element to vibrate at its mechanical resonance (tuned to 40,000 Hz) using the principles of magnetostriction. The shift in the measured frequency, which corresponds to an ice thickness of approximately 0.51 mm (0.02 in.) triggers the ICE DETECTED discrete. When ice is detected, the strut and probe are de-iced and melt ice accumulation. The aerodynamic airflow then sheds the ice and provides the probe cooling for continued sensing. When it is powered up and operational, the ice detector provides an ICE DETECTOR OK signal. When icing is detected, each detector provides three signals: ICE DETECTED 1, ICE DETECTED 2, and ICE DETECTED 3. The ICE DETECTED 1 and ICE DETECTED 2 signals are sent to the DMCs and distributed to the left and right electronic engine controls (EECs) for cowl anti-ice (CAI) and to the three primary flight computers (PFCCs). The DMC also sends the ice detection signals to the integrated air system controllers (IASCs) to enable wing anti-ice system (WAIS) activation. ICE DETECTED 3 signal is sent through the control and distribution cabinets (CDCs) for the WAIS. The ICE DETECTED and the ICE DETECTOR OK signals are used to generate caution and advisory EICAS messages. The ice detection system is also monitored by the OMS. An ICE caution message is displayed when ice is detected and the wing anti-ice system (WAIS) or cowl anti-ice system (CAIS) are inoperative. An ICE advisory message indicates normal operation of wing and cowl anti-ice systems in icing conditions. When the ICE advisory message is removed the crew can select the anti-ice systems off. If an ice detector fails, the associated ICE DET FAIL caution message is displayed. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-6
CS130_30_L3_IDS_DetailedDescription.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
CDC 3-14-5
CDC 4-13-4 L ICE DET
AC BUS 1
R ICE DET AC BUS 2
SSPC 5A
Logic: Always On
SSPC 5A
Logic: Always On
LEFT ICE DETECTOR
L CAIS
R CAIS
Test
Test L EEC
R EEC
Ice Detector OK
Ice Detector OK
EICAS OMS
Ice Detected 1 Ice Detected 2
RIGHT ICE DETECTOR
DMC 1
Ice Detected 1 DMC 2
Ice Detected 2
PFCCs
CAIS CDC DMC EEC IASC PFCC WAIS
Cowl Anti-Ice System Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Engine Control Integrated Air System Controller Primary Flight Control Computer Wing Anti-Ice System AFDX ARINC 429 Discrete TTP BUS
IASC 1
Ice Detected 3
WAIS
IASC 2
CDC 1
CDC 2
CDC 3
CDC 4
Ice Detected 3
CS1_CS3_3081_014
LEGEND
Figure 3: Ice Detection System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-7
CS130_30_L3_IDS_DetailedDescription.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the ice detection system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-8
CS130_30_L2_IDS_MonitoringTests.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L ICE DET FAIL
Left ice detector has failed.
R ICE DET FAIL
Right ice detector has failed.
ICE
Ice detected and wing anti-ice or cowl anti-ice are failed or is off.
Table 2: ADVISORY Messages MESSAGE ICE
LOGIC Ice detected but ice detection, wing anti-ice and cowl anti-ice are working properly.
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30-9
CS130_30_L2_IDS_MonitoringTests.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
INITIATED BUILT-IN TEST Once powered, each ice detector initiates its own power-up built-in test (PBIT) lasting about 7 seconds. The detector produces an ICE DETECTOR OK signal after internal testing is successful and remains active, provided it continues to function normally. An initiated built-in test (IBIT) function performs a diagnostic check of both ice detectors, including communication with interconnecting systems. The IBIT is initiated from the EICAS AVIONICS page on the AVIO tab. When IBIT is initiated: •
A message IN PROG is provided
•
The ice detectors perform their internal tests
•
The data concentrator unit module cabinets (DMCs) monitor ice detection signals, ICE DETECTED 1, and ICE DETECTED 2, from the ice detectors
•
The DMCs monitor ice detection signal, ICE DETECTED 3, from the control and distribution cabinets (CDCs)
When the ice detectors internal test is complete, the detector produces an ICE DETECTOR OK signal if internal testing is successful. One of two messages, PASS or FAIL, is displayed on the AVIONICS synoptic page on the AVIO tab and more data is available in the onboard maintenance system (OMS) The PASS or FAIL message is triggered by the value of ICE DETECTOR OK signal. At any time, if an ice detector has an internal fault, the ice detector turns off and a L (R) ICE DET FAIL caution message is displayed.
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30-10
CS130_30_L2_IDS_InitiatedBIT.fm
30 - Ice and Rain Protection 30-81 Ice Detection System
SYNOPTIC PAGE - AVIONIC
LEGEND CDC DMC
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet ARINC 429 Discrete
STATUS
AIR IR
FUEL FUE EL
ELEC C
HYD YD
INFO FO
CBB
CTP CT TP CTP
AVIO
TEST STATUS
FLT CTRL CTRL
SMS RUNWAY
ENABLED
Status
Condition
PASS
Test successfully completed
FAULT
Test failure with fault message
FAIL
Test failure
IN PROG
Test in progress
----
Test invalid or aborted
VSPEEDS...
TEST
LEFT ICE DETECTOR
AURAL AURA URA AL
WING ING A/IC A/I A/ICE CE
LAMP P
PASS ICE DETECT
TCAS CAS
RIGHT ICE DETECTOR
FIRE FIR E
Test
Test
WXR
FLT CTRL CTRL
TAWS S
SHAKER R
Ice Detected 1 and Detected Ice 2
Ice Detector OK DMC 1
Ice Detected 3
Test
Test
CDCs
DMC 2
Ice Detected 1 and Detected Ice 2
Ice Detected 3
CS1_CS3_3081_015
Ice Detector OK
Figure 4: Ice Detection System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-11
CS130_30_L2_IDS_InitiatedBIT.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
30-12 WING ANTI-ICE SYSTEM GENERAL DESCRIPTION The wing is protected from icing conditions by the wing anti-ice system (WAIS), which prevents the formation of ice on the leading edge of wing slats no. 2, no. 3, and no. 4 by supplying the hot bleed air from the engines. Ice protection is not provided for the slats that are located inboard of the engines. The wing anti-ice valves (WAIVs) regulate the pressure to 27 psi in the associated wing anti-ice ducting. Flow is controlled by venturis. After heating the inside of the slat, the bleed air is expelled overboard underneath the wing. The WAIS includes two independent anti-icing systems for the left and right wings. Each wing system is connected to the other by a normally closed cross-bleed valve (CBV). The engine bleed air system supplies the hot air at required pressure and temperature to the distribution ducts installed in the slats. If one engine or bleed air supply fails, the bleed air system opens the CBV to supply air to the opposite wing. The integrated air system controllers (IASCs) send a wing anti-ice (WAI) enable signal to the data concentrator unit module cabinets (DMCs) to activate or deactivate the wing anti-ice system (WAIS), based on the ice detection system and WING switch position.
When ANTI-ICE WING switch is set to ON, the WAIS is activated regardless of icing conditions if the airspeed is above 60 kt on the ground or anytime in flight. The WAIS can be tested on the ground if the switch is in the ON position and the airspeed is below 60 kt. The test lasts 20 seconds. Each wing is continuously monitored for temperature and pressure by both IASCs. The wing anti-ice temperature sensor (WAITS) provides monitoring of the temperature of the bleed air supplied to the wing ducting. The pressure sensor measures system pressure from the outboard end of slat no. 4. Pressure is monitored to detect excessive high pressure that might lead to ducting failures, as well as low pressure which may indicate inadequate hot air for effective anti-icing or to detect a duct burst or leak. In case a leak is detected in the ducting network, the IASC shuts down the WAIS automatically. IASC 1 is powered by DC ESS BUS 1 and DC ESS BUS 2. IASC 2 is powered by DC ESS BUS 2 and DC ESS BUS 3. The WAIVs are powered by DC ESS BUS 2 provided by the control and distribution cabinets (CDCs). The pressure sensors are powered from DC ESS BUS 3.
The WAIS is monitored and controlled by the IASCs. With the ANTI-ICE WING switch in the AUTO position, the IASCs automatically activate the WAIS on receiving ice signals from one or both ice detectors. The WAIS is monitored and controlled by the IASCs. In AUTO mode, the WAIS is inhibited on the ground and during takeoff. In flight, the IASCs automatically activate the WAIS on receiving ice signals from one or both ice detectors.
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30-12
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
ICE DETECTION SYSTEM
NOTE IASC 1 is powered by DC ESS BUS 1 and DC ESS BUS 2. IASC 2 is powered by DC ESS BUS 2 and DC ESS BUS 3.
CDCs (DC ESS BUS 2) ANTI-ICE WING
EICAS
AUTO
OFF
ON
DMCs
SYN OMS
ANTI-ICE PANEL IASCs
LEGEND
PS SOL
Pressure Sensor
CBV SOL
SOL
Slats
Slats WAIV
WAIV
WAITS
WAITS
PS
Venturi
Solenoid Temperature Sensor
DC ESS BUS 3
PS
CS1_CS3_3012_019
CBV CDC DMC IASC WAITS WAIV
Bleed Air Wing Anti-Ice Air Cross-Bleed Valve Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve ARINC 429 Discrete
Figure 5: Wing Anti-Ice System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-13
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE DUCTING The wing anti-ice ducting consists of a venturi duct and three piccolo tubes connected by interslat hoses. The air exits the piccolo tubes and circulates around the front bay inner chamber of the slat to be expelled overboard through slots and exhaust holes in the slat rear bay. A telescopic duct connects the wing anti-ice system air duct to the piccolo tubes to allow for the slats extension and retraction.
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30-14
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Engine Bleed Air Supply
Telescopic Duct Interslat Hoses
Venturi
SOL
WAIV
WAITS
Rear Bay Seal
Front Bay
PS
Piccolo Tubes
Exhaust Holes
Piccolo Tube
WAITS WAIV
Bleed Air Wing Anti-Ice Air Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve
PS
Pressure Sensor
SOL
Solenoid
Slat
CROSS-SECTION VIEW
CS1_CS3_3012_012
LEGEND
Figure 6: Wing Anti-Ice Air Supply Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-15
CS130_30_L2_WAIS_General.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
COMPONENT LOCATION
PRESSURE SENSORS
The wing anti-ice system (WAIS) consists of the following components:
Each pressure sensor is connected by a tube, to the outboard end of the piccolo tube assembly in slat no. 4.
•
Wing anti-ice valves (WAIVs)
•
Venturis
•
Telescopic ducts
•
Piccolo tubes
•
Wing anti-ice temperature sensors (WAITSs)
•
Pressure sensors (PSs)
•
Integrated air system controllers (IASCs) (Refer to figure 8).
WING ANTI-ICE VALVES Each wing anti-ice valve (WAIV) is located in the wing leading edge, outboard of the engine pylon, forward of the front spar. VENTURIS A venturi is located immediately downstream and outboard of the WAIV, in the wing leading edge cavity aft of slat no. 2. TELESCOPIC DUCTS Each telescopic duct is located in the wing leading edge cavity aft of slat no. 2. PICCOLO TUBES On each wing, slat no. 2, no. 3, and no. 4 each contain a piccolo tube. The piccolo tubes are joined by interslat hoses. WING ANTI-ICE TEMPERATURE SENSORS Each wing anti-ice temperature sensor (WAITS) is installed downstream of the venturi duct, before the telescopic duct.
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30-16
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
B
A Slat No. 2
Wing Anti-Ice Valves (2x)
A Slat No. 3
C Slat No. 4
Venturis (2x)
Telescopic Ducts (2x)
Wing Anti-Ice Temperature Sensor
A Piccolo Tubes (6x)
Interslat Hose
Pressure Sensors (2x)
NOTE Left wing shown. Right wing similar. Slats not shown for clarity.
C
CS1_CS3_3012_013
B
Figure 7: Wing Anti-Ice System Leading Edge Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-17
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
INTEGRATED AIR SYSTEM CONTROLLERS There are two integrated air system controllers (IASCs) located in the mid equipment bay.
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30-18
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
IASC 2
IASC 1
INTEGRATED AIR SYSTEM CONTROLLERS (2X)
CS1_CS3_3012_017
MID EQUIPMENT BAY FWD RACK
Figure 8: Integrated Air System Controllers Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-19
CS130_30_L2_WAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
CONTROLS AND INDICATIONS ANTI-ICE CONTROL PANEL The wing anti-ice system (WAIS) has a three-position rotary switch located on the overhead panel of the flight deck. The switch is normally left in the AUTO position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-20
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
A
OVERHEAD PANEL
ANTI-ICE L COWL
WING
AUTO
OFF
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
CS1_CS3_3012_021
ANTI-ICE PANEL
Figure 9: Wing Anti-Ice Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-21
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE INDICATIONS The wing anti-ice system (WAIS) indications and messages are displayed on the EICAS page and the AIR synoptic page. EICAS Page The wing anti-ice (WAI) icon is displayed on EICAS when the system is not OFF. The WAI icon is displayed in green when the system is activated, displayed in white when inhibited, or amber when a failure is reported. AIR Synoptic Page The status of the wing anti-ice valves (WAIVs) and the wing leading edge are displayed on the AIR synoptic page.
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30-22
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
CLB
STATUS TUS
92.2
92.2
73.0
73.3
FUEL FUE EL
722
EGT N2 FF (KPH) OIL TEMP OIL PRESS
ELECC
WING ANTI-ICE VALVE
FLT CTR CTRLL
HYD YD
AVIONIC NIC
INFO
CB
Condition Open with flow Closed
718
86.5 1090 124 116
DOOR OR
Symbol
N1
CAI WAI
AIR IR
86.6 1110 124 116
LO
FWD 22 °C 22 °C -+ 2
AFT 22 °C 22 °C -+ 2
10 °C °
15 °CC
15 °CC
CKPT 23 °C 22 °C
CARGO
15°CC
Open with no flow (normal) Open with no flow (fail) Open with flow (fail)
CAI WAI
RAM AIR
TRIM AIR L PACK
HI
R PACK
Invalid Closed (fail)
HI
EICAS WING ANTI-ICE STATE
45
WING ANTI-ICE SYSTEM Symbol
Condition
WAI
Wing anti-ice activated
WAI
Wing anti-ice inhibited
WAI
Wing anti-ice caution failure or overheat
WAI
WAI failed in warning condition
PSI
XBLEED
45 PSI
Symbol
Condition Off Normal
Left Wing Anti-Ice Valve
Abnormal
Right Wing Anti-Ice Valve
SYNOPTIC PAGE – AIR
CS1_CS3_3012_001
APU
Figure 10: Wing Anti-Ice Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-23
CS130_30_L2_WAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
DETAILED DESCRIPTION INTEGRATED AIR SYSTEM CONTROLLERS The integrated air system controllers (IASCs) control and monitor wing anti-ice system (WAIS) operation using position sensing of the wing ant-ice valve (WAIV) (full closed), wing anti-ice (WAI) temperature and pressure. Each IASC has three channels: channel A, channel B, and a safety channel. Channel A is not used by the WAIS. The WAIS uses the IASC to: •
Inhibit or enable activation
•
Monitor WAIS pressure and temperature
•
Monitor WAIS component health
When the aircraft is powered, the active IASC is determined from odd dates for IASC 1 and even dates for IASC 2. Channel B of the active IASC for the day is assigned to WAIS activation. The safety channels of both IASCs, in conjunction with channel B of the standby IASC, monitor WAIS temperature and pressure and provide signals to the EICAS, through the data concentrator unit module cabinets (DMCs). The standby IASC also monitors WAIS component health by performing a continuous built-in test (CBIT). Data from CBIT is reported to the onboard maintenance system (OMS) through the DMCs The operation of the WAIS is shown schematically on the AIR synoptic page. Synoptic page data comes from the safety channel through the DMCs.
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CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Channel A
Channel B
Safety Channel
L/R WAIS Activation
L/R WAIS Monitoring: • Pressure • Temperature L/R Component Health Monitoring
IASC (Standby) L PRESSURE SENSOR
L WAITS
L WAIV
EICAS DMC SYN
R PRESSURE SENSOR
R WAITS
OMS
R WAIV
L/R WAIS Monitoring: • Pressure • Temperature L/R Component Health Monitoring
Channel A
Channel B
Safety Channel
IASC (Active)
LEGEND DMC IASC OMS WAIS WAITS WAIV
Data Concentrator Unit Module Cabinet Integrated Air System Controller Onboard Maintenance System Wing Anti-Ice System Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve ARINC 429
CS1_CS3_3012_010
L/R WAIS Activation
Figure 11: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-25
CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE OPERATION The wing anti-ice system (WAIS) uses precooled engine bleed air, sprayed by the piccolo tubes onto the internal surface of the wing leading edge slats to prevent ice buildup. The bleed air system controls engine bleed air supply pressure and temperature to the WAIS. The supply bleed air temperature target at the bleed temperature sensor (BTS) is nominally 227°C (440°F) in flight. On the ground, during initiated built-in test (IBIT), the temperature target is 200°C (392°F). The WAIV controls the pressure of bleed air to the wing leading edge slats. Wing anti-ice temperature sensors (WAITSs) are used by the integrated air system controller (IASC) to monitor the bleed air temperature, downstream of the venturis. Downstream of the wing anti-ice valve (WAIV), the hot bleed air passes through the venturi and telescopic duct to reach the slat no. 2 piccolo tube, then to slat no. 3 and no. 4 through interslat hoses. The area between the slats is not heated.
The pressure sensor detects pressure at the end of the piccolo tubes of slats no. 4. A low-pressure condition of less than 14 psi for more than 15 seconds generates a WING ANTI-ICE FAIL warning message on EICAS for the associated wing. If the low-pressure condition persists and results in a low-temperature condition on the leading edge, the WING ANTI-ICE LO HEAT caution message replaces the previous warning message. If an overpressure condition of more than 21 psi for more than 15 seconds (indicative of a WAIV failed open condition) is detected by the pressure sensor (PS), a WING ANTI-ICE FAIL warning message is displayed on EICAS for the associated wing. This message also appears if the WAIV if not closed when on the ground. If the WAIV position does not indicate full closed with the WAIS off, a WING ANTI-ICE FAIL ON caution message is displayed on EICAS for the associated wing.
The purpose of the venturi is to choke the air flow in case of a burst or crack in the downstream ducting. The dual-channel WAITS monitors temperature downstream of the WAIV. If a high-temperature condition is detected that exceeds 254°C (489°F) for more than 30 seconds, a WING ANTI-ICE OVHT caution message is displayed on EICAS for the associated wing. If a low-temperature condition of less than 200°C (392°F) is detected by the WAITS for more than 60 seconds, a WING ANTI-ICE LO HEAT caution message is displayed on EICAS for the associated wing. The actual low-temperature threshold is a function of aircraft altitude and pressure reported by the pressure sensor.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
ICE DETECTION SYSTEM
1 CDC 2-5-17 WING A-ICE VLV DC ESS SSPC 3A BUS 2 Logic: Combinational CDC 2
Ice Detected 3
CDC 3 and CDC 4
CDC 2
1 ANTI-ICE WING
Ice Detected 1
OFF
Ice Detected 2
EICAS
AUTO
ON
DMCs
SYN ANTI-ICE PANEL OMS
IASC 1
IASC 2
LEGEND
PS SOL
Pressure Sensor
Solenoid
CBV
SOL
Slats
SOL
Slats
WAIV
WAIV
WAITS
WAITS BTS CB-D2 PS
DC ESS BUS 3
Venturi ECS SAFETY SNSR
Temperature Sensor
PS
CS1_CS3_3012_011
BTS CBV CDC DMC IASC WAITS WAIV
Bleed Air Wing Anti-Ice Air Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Integrated Air System Controller Wing Anti-Ice Temperature Sensor Wing Anti-Ice Valve AFDX ARINC 429 Discrete TTP BUS
Figure 12: Wing Anti-Ice System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-27
CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
WING ANTI-ICE ACTIVATION CONTROL The WAIVs are commanded to simultaneously open when they receive 28 VDC from control and distribution cabinet 2 (CDC 2) and controlled by advanced logic. The combinational logic within CDC 2 receives input from the following sources: •
An enable/inhibit signal from the IASCs via the DMCs
•
The WING anti-ice switch
•
Ice detection signals from the ice detection system
•
Airspeed < 60 kt and WAIS test selected
-
Airspeed > 60 kt Anytime the switch is in ON position
With WAIS switch in AUTO position, and ice is detected, the system turns ON: •
Main engine start (MES)
•
Bleed air leak detected by bleed air leak and overheat detection system (BALODS)
•
Wing anti-ice (WAI) switch selected OFF
When in flight, the WAIS can be safely operated below 12°C (53°F) total air temperature (TAT) or airspeed above 60 kt. A WING A/ICE ON caution message is provided by the system when the system is ON and the TAT is above 15°C (59°F).
On ground: -
•
•
Any loss of redundant or non-critical function results in a WING A/ICE FAULT advisory message.
In flight: -
No engine bleed is available (one engine minimum required)
When the WAIS rotary switch is set to ON or AUTO and neither engine is running, a WING A/ICE MISCONFIG caution message is displayed on EICAS.
On ground: -
•
The WAIVs are closed 2 minutes after the ICE DETECT signal is no longer present.
With the WAIS switch in the ON position, the system turns ON: •
The WAIS is automatically deactivated when any of the following occurs:
Never
In flight, in either of the following conditions: -
Altitude > departure airport altitude + 457 m (1,500 ft)
-
Takeoff and 2 minutes after airspeed > 60 kt
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
AUTO MODE
MANUAL MODE
Engine Bleed Air ACFT on Ground Airspeed < 60 kt WAIS Test Enabled WAIS Switch ON
Takeoff Airspeed > 60 kt + 2 Minutes
20 Second Timer
ALT. > Departure Airport ALT. + 1.500 ft WAIS Valve Open
WAIS Valve Open
Aircraft in Flight Ice Detected WAIS Switch in AUTO
CS1_CS3_3012_020
ACFT in Flight
LEGEND WAIS
Wing Anti-Ice Switch
Figure 13: Wing Anti-Ice Activation Advanced Logic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-29
CS130_30_L3_WAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
MONITORING AND TESTS SYSTEM TEST A initiated built-in test (IBIT) is available on the AVIONICS synoptic page on the AVIO tab. The test can be carried out on ground, with the bleed air system pressurized by the engines. IBIT is limited to one cycle prior to takeoff to protect the leading edges from damage caused by excessive heat cycles. The airplane flight manual (AFM) limits the test to once per day. In case of test failure, the AFM limits the time between tests to a minimum of 30 minutes. The WING anti-ice switch must be set to the ON position for the test. If the auxiliary power unit (APU) bleed air is being used, the IASCs automatically switch to engine bleed. The icon is grayed out if conditions are not correct for the IBIT. When the WING A/ICE icon is pressed, the IASCs open the WAIVs and monitor the pressures obtained. The test period is limited to 20 seconds, after which the WAIVs close. The initiated built-in test (IBIT) results are displayed on the synoptic page as either PASS, FAIL, or FAULT. IBIT does not activate if a failure is detected and not rectified, indicated by the FAIL message. After failure rectification, the test can be re-initiated by resetting the IASCs. A wait time of 30 minutes is required to ensure cooling of the wing slats before repeating the test. A display of ----- indicates the system had communication problems during the test. Additionally, the IASCs conduct a power-up built-in test (PBIT) when power is first available on the aircraft. This check determines the validity of the two pressure sensor signals.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-30
CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
TEST STATUS Status
Condition
PASS
Test successfully completed
FAULT
Test failure with fault message
FAIL
Test failure
IN PROG
Test in progress
----
Test invalid or aborted
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
CTP
AVIO SMS RUNWAY
ENABLED VSPEEDS...
Engines Running On Ground
TEST
AURAL
PASS WING A/ICE
LAMP
ICE DETECT
TCAS
FIRE
WXR
FLT CTRL
TAWS
SHAKER
L COWL
WING
AUTO
OFF
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
CS1_CS3_3012_005
ANTI-ICE
Figure 14: Wing Anti-Ice System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-31
CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
CAS MESSAGES The following page provides the crew alerting system (CAS) and INFO messages associated with the wing anti-ice system (WAIS).
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Table 6: STATUS Messages Table 3: WARNING Messages MESSAGE
MESSAGE
LOGIC
L WING A/ICE FAIL
L WAI not closed in any condition on ground out of test sequence, overpressure or low-pressure.
R WING A/ICE FAIL
R WAI not closed in any condition on ground out of test sequence, overpressure or low-pressure.
WING A/ICE ON
Wing anti-ice selected ON.
WING A/ICE OFF
Wing anti-ice selected OFF.
Table 7: INFO Messages MESSAGE 30 WING A/ICE FAULTWING A/ICE TEMP SNSR REDUND LOSS
Table 4: CAUTION Messages MESSAGE
LOGIC
LOGIC
LOGIC Single failure of one WAITS channel.
L WING A/ICE OVHT
Left wing bleed overtemperature detected.
R WING A/ICE OVHT
Right wing bleed overtemperature detected.
30 L WING A/ICE LO HEAT L WAITS drift low. - L WING A/ICE TEMP SNSR INOP
L WING A/ICE LO HEAT
Low combination of pressure and temperature for L WAI (also function of altitude).
30 L WING A/ICE LO HEAT L bleed temperature control too low for WAI. - CTRL TEMP INOP
R WING A/ICE LO HEAT
Low combination of pressure and temperature for R WAI (also function of altitude).
30 L WING A/ICE LO HEAT L HPV failed closed leading to WAI low - L HPV FAIL CLSD temperature.
WING A/ICE FAIL
Left or right WAI function failed.
WING A/ICE ON
Wing anti-ice selected ON and TAT above 15°C (59° F).
30 L WING A/ICE OVHT - L L WAITS drift high. WING A/ICE TEMP SNSR INOP
WING A/ICE MISCONFIG
WAI selected ON or WAI AUTO and ice detected with both engine bleeds OFF.
30 L WING A/ICE OVHT CTRL TEMP INOP
L bleed temperature control too high for WAI.
30 R WING A/ICE LO HEAT - R WING A/ICE TEMP SNSR INOP
R WAITS drift low.
Table 5: ADVISORY Messages MESSAGE WING A/ICE FAULT
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC Loss of redundant or non-critical function for the wing anti-ice system or WAIV high leakage. Refer to INFO messages.
30 R WING A/ICE LO R bleed temperature control too low for WAI. HEAT - CTRL TEMP INOP 30 R WING A/ICE LO HEAT - R HPV FAIL CLSD
R HPV failed closed leading to WAI low temperature.
30 R WING A/ICE OVHT R WING A/ICE TEMP SNSR INOP
R WAITS drift high.
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Page Intentionally Left Blank
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Table 7: INFO Messages MESSAGE
LOGIC
30 R WING A/ICE LO R bleed temperature control too low for WAI. HEAT - CTRL TEMP INOP 30 WING A/ICE FAULT Loss of ICE DETECTED signal at IASC input. WING A/ICE AUTO MODE INOP 30 WING A/ICE FAULT - L Erroneous L WAIV switch acquisition on one WING A/ICE VLV POS SW channel. INOP 30 WING A/ICE FAULT - R Erroneous R WAIV switch acquisition on one WING A/ICE VLV POS SW channel. INOP 30 WING A/ICE FAIL - L WING PRESS FAIL
Left wing anti-ice overpressure or low-pressure.
30 WING A/ICE FAIL - R WING PRESS FAIL
Right wing anti-ice overpressure or low-pressure.
30 WING A/ICE FAULT - L WING A/ICE VLV LEAK
Left WAIV internal leakage.
30 R WING A/ICE FAIL - R Right WAIV internal leakage. WING A/ICE VLV LEAK 30 L WING A/ICE FAIL - L WING A/ICE VLV FAIL OPEN
Left wing anti-ice failed open.
30 R WING A/ICE FAIL - R Right wing anti-ice failed open. WING A/ICE VLV FAIL OPEN 30 WING A/ICE FAULT - L WING A/ICE PRESS SNSR INOP
Left outboard pressure sensor out of range or drifted.
30 R WING A/ICE FAIL - R Right outboard pressure sensor out of range or WING A/ICE VLV LEAK drifted.
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CS130_30_L2_WAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
PRACTICAL ASPECTS WING ANTI-ICE VALVE For dispatch purposes, the wing anti-ice valve (WAIV) is equipped with a manual override provision to rotate the valve to the closed position. A deactivation screw locks the valve in the closed position. When the deactivation screw is removed and used to secure the valve closed, air is bled from the pneumatic actuator line to prevent the valve from operating.
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30-36
CS130_30_L2_WAIS_PracticalAspects.fm
30 - Ice and Rain Protection 30-12 Wing Anti-Ice System
Deactivation Screw in Stowed Position
Deactivation Screw Position for Valve Lockout
CS1_CS3_3011_003
Manual Override
Figure 15: Wing Anti-Ice Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-37
CS130_30_L2_WAIS_PracticalAspects.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
30-22 COWL ANTI-ICE SYSTEM GENERAL DESCRIPTION The cowl anti-ice system (CAIS) uses engine bleed air from the 6th stage compressor. The air is supplied to the engine inlet cowl, heating the leading edge and preventing ice accumulation.
The EEC activates the system based on switch settings from the ANTI-ICE panel and the ice detection system signals. The CAIS operates when the EEC commands a pair of CAIVs to open.
The cowl anti-ice valves (CAIVs) provide a pressure regulating and shutoff function. The two valves are similar in design, and provide redundancy in case of the failure of a single valve, however, the CAIVs are not interchangeable and have different part numbers.
A dual-element fan cowl temperature sensor in the fan cowl sends fan case area temperature to the EEC. This sensor is used to detect a temperature rise due to a possible CAIS duct rupture.
Each CAIV is a spring-loaded, failsafe open, poppet-type valve. Closure of the valve requires electrical power applied to the solenoid assembly and bleed air pressure. The valves are open when the solenoid assembly is electrically de-energized and/or there is a loss of bleed air pressure. Each engine CAIS operates independently. The COWL rotary switches located on the ANTI-ICE panel, activates the CAIS when set to the AUTO position, or when selected to the ON position in manual mode. Primary power for both electronic engine controls (EECs) is from their respective engine-driven permanent magnet generator (PMG) supplying 3-phase 34 VAC. Each EEC channel automatically switches power source to aircraft power source should the PMG fail. The left EEC is powered by DC ESS BUS 3 for channel A, and DC ESS BUS 1 for channel B. The right EEC is powered by DC ESS BUS 3 for channel A, and DC ESS BUS 2 for channel B.
A dual-channel pressure sensor sends CAIS duct pressure to the EEC for monitoring CAIS operation. When the engine reaches ground idle after starting, the EEC commands one of the CAIVs to close and monitors downstream bleed pressure, confirming its ON/OFF capability. Each CAIV is checked alternately every other flight. System messages are sent by the EEC to the engine indication and crew alerting system (EICAS), and the onboard maintenance system (OMS) through the data concentrator unit module cabinets (DMCs). System operation is displayed schematically on the AIR synoptic page.
The operational logic for the CAIS is contained in the EEC, operating on two channels, A and B. Further details on the EEC can be found in ATA 73 Engine Fuel and Control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_30_L2_CAIS_General.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
ANTI-ICE L COWL
WING
AUTO
OFF
Fan Cowl Temperature Sensor
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
Fan Cowl Temperature Sensor
ANTI-ICE PANEL EICAS
Pressure Sensor
Pressure Sensor
SYN OMS 6th Stage Air
PMG
PMG PS
PS
DMCs
EEC
EEC
Ch B
Ch A
ICE DETECTION SYSTEM
EEC
EEC
Ch A
Ch B
6th Stage Air
Cowl Anti-Ice Valve (2x)
Cowl Anti-Ice Valve (2x) DC ESS BUS 2 DC ESS BUS 3 DC ESS BUS 1
Bleed Air Data Concentrator Unit Module Cabinet Electronic Engine Control Permanent Magnet Generator (3-Phase 34 VAC) ARINC 429 Discrete
DC ESS BUS 3 CS1_CS3_3021_006
DMC EEC PMG
LEGEND
Figure 16: Cowl Anti-Ice System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-39
CS130_30_L2_CAIS_General.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
COMPONENT LOCATION The cowl anti-ice system (CAIS) consists of left and right subsystems each of which contain: •
Cowl anti-ice valves (CAIVs)
•
Pressure sensor
•
Fan cowl temperature sensor
•
Swirl nozzle
COWL ANTI-ICE VALVES The two cowl anti-ice valves (CAIVs) are mounted one above the other at the 6 o’clock position under the engine compressor case. PRESSURE SENSOR The pressure sensor is located on the engine fan case aft section, at the 5 o’clock position. FAN COWL TEMPERATURE SENSOR The fan cowl temperature sensor is located on the right side of the engine fan case at the 4:30 o’clock position. SWIRL NOZZLE The swirl nozzle duct is located within the engine inlet cowl.
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CS130_30_L2_CAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
A COWL ANTI-ICE VALVES (2X)
B PRESSURE SENSOR
A B
C
C FAN COWL TEMPERATURE SENSOR
D
SWIRL NOZZLE
CS1_CS3_3021_007
D
Figure 17: Cowl Anti-Ice System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-41
CS130_30_L2_CAIS_ComponentLocation.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
DETAILED COMPONENT INFORMATION SWIRL NOZZLE The swirl nozzle has an inner wall and an outer wall. Hot air exiting the nozzle swirls within the inlet cowl, heating the surface of the inlet cowl lip to provide anti-icing. In the event of a rupture to the inner wall, the outer wall contains the hot engine bleed air directing it backwards toward an insulation boot. The insulation boot is designed to allow hot air from a ruptured inner wall to blow back into the engine fan casing area. The rupture is detected by the fan cowl temperature sensors.
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30-42
CS130_30_L3_CAIS_DetailedComponentInfo.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
Inlet Cowl Aft Bulkhead
Insulation Boot (deforms when hot) Air Flow
Hot Air Leak
Exhaust Slots Inner Wall
SWIRL NOZZLE WITH LEAK
Outer Wall
CS1_CS3_3021_008
Fan Cowl Temperature Sensor
Figure 18: Swirl Nozzle Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-43
CS130_30_L3_CAIS_DetailedComponentInfo.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
CONTROLS AND INDICATIONS The cowl anti-ice system (CAIS) is controlled by two switches located on the anti-ice panel to provide separate controls for the left, and right systems.
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CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
A
OVERHEAD PANEL
ANTI-ICE L COWL
WING
AUTO
OFF
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
CS1_CS3_3021_009
ANTI-ICE PANEL
Figure 19: Cowl Anti-Ice System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-45
CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
INDICATIONS Indication of cowl anti-ice system (CAIS) operation is displayed on EICAS and the AIR synoptic page. The cowl anti-ice (CAI) icon is displayed on EICAS to advise the crew of the system status. The CAI icon on EICAS is displayed in GREEN when the system is activated, displayed in WHITE when inhibited, or AMBER when a failure is reported.
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30-46
CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
CLB
STATUS TUS
92.2
92.2
73.0
73.3
718
722
EGT
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
DOOR OR
ELECC
FLT CTR CTRLL
COWL ANTI-ICE VALVE Symbol
FUEL FUE EL
HYD YD
AVIONIC NIC
INFO
CB
Condition Open with flow Closed
N1
CAI WAI
AIR IR
86.6 1110 124 116
LO
FWD 22 °C 22 °C -+ 2
AFT 22 °C 22 °C -+ 2
10 °C °
15 °CC
15 °CC
CKPT 23 °C 22 °C
CARGO
15°CC
Open with no flow (normal) Open with no flow (fail) Open with flow (fail)
CAI WAI
RAM AIR
TRIM AIR L PACK
HI
R PACK
Invalid Closed (fail)
HI
EICAS COWL ANTI-ICE STATE COWL ANTI-ICE ICON Symbol
45 PSI
XBLEED
45 PSI
Condition
Symbol
Condition Off
CAI
Cowl anti-ice activated
Normal
CAI
Cowl anti-ice failure reported
Abnormal
CAI
Cowl anti-ice inhibited
SYNOPTIC PAGE – AIR
CS1_CS3_3022_001
APU
Figure 20: Cowl Anti-Ice System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30-47
CS130_30_L2_CAIS_ControlsIndications.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
DETAILED DESCRIPTION
TEMPERATURE AND PRESSURE SENSING
The cowl anti-ice system (CAIS) uses bleed air supplied from the 6th stage of the high-pressure compressor (HPC) to heat the engine inlet cowl lip.
The fan cowl inlet temperature sensor and the cowl anti-ice (CAI) pressure sensor on each engine send signals to EEC for monitoring.
When the CAIS is in operation, the cowl anti-ice valves (CAIVs) are commanded open allowing hot air flow to the inlet cowl. The hot bleed air swirls through the inlet cowl and is discharged overboard through the exhaust louvers. If one of the two CAIVs remains closed, a L (R) COWL A/ICE FAIL caution message on the EICAS appears. In flight, when the COWL switch is selected to ON, the CAIS is activated regardless of ice detection. If the outside air temperature (OAT) is greater than 15°C (60°F), the COWL ANTI-ICE ON caution message is displayed on EICAS for the associated engine. On the ground, the CAIS is inhibited when selected to ON if the outside air temperature (OAT) is above 15°C (60°F). In flight, when the COWL switch is selected to AUTO, the CAIS is activated when ice is detected. The CAIS is turned off when the ice detectors no longer signal ice (2 minutes after leaving icing conditions). If the ice signals are not valid, the electronic engine control (EEC) assumes icing conditions and opens the CAIVs. On the ground in AUTO mode, CAIS is inhibited with takeoff thrust commanded. The EEC maintains the takeoff inhibit condition until either the aircraft has climbed 457 m (1500 ft) above departure airport elevation, or 2 minutes have elapsed since takeoff inhibit was set.
In case of a rupture to the CAIS ducting, the affected temperature sensor sends signals to the EEC. The EEC generates an ENG NACELLE OVHT caution message in the EICAS only if the EEC is not able to close the CAIV. With the CAIV commanded closed due to overheat, the EEC continues to command it closed until the EEC is reset. Unless reset, once the CAIS system is required to provide anti-ice heat in ON or AUTO mode, the EEC generates a COWL A/ICE FAIL caution message on EICAS. While in operation, if the CAIS pressure falls below 40 psi, the pressure sensor signal causes the EEC to generate a COWL A/ICE FAIL caution message. If CAIS pressure rises above 71 psi, an ENGINE FAULT advisory message is displayed. Any time an engine is running, and the EEC does not detect valid ICE DETECTION signal, the EEC opens the CAIS valves. During an engine start, both cowl anti-ice valves (CAIVs) are open until the engine reaches ground idle. Once the engine reaches ground idle, the electronic engine control (EEC) commands one of the CAIVs closed. This enables the EEC to verify system control. Each valve is checked alternately on every start. After the valve is checked, the EEC commands both valves to close.
The EEC provides system data through the data concentrator unit module cabinets (DMCs) for EICAS messages, the AIR synoptic page and for the onboard maintenance system (OMS).
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CS130_30_L3_CAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
LEGEND CAIS DMC EEC HPC
Bleed Air Cowl Anti-Ice System Data Concentrator Unit Module Cabinet Electronic Engine Control High-Pressure Compressor ARINC 429 Discrete
ICE DETECTION SYSTEM
ANTI-ICE L COWL
EICAS
WING
AUTO
OFF
SYN
R COWL
AUTO
ON
OFF
AUTO
ON
OFF
ON
DMCs
OMS ANTI-ICE PANEL
DC ESS BUS 1 DC ESS BUS 3
AUTO MODE Takeoff Airspeed > 60 kt + 2 Minutes ALT. > Departure Airport ALT. + 1.500 ft
EEC
Aircraft in Flight Ice Detected
CAIS Valve Open
CAIS Switch in AUTO
Fan Cowl Temperature Sensor
Aircraft On Ground Outside Air Temperature < 15°C (60°F)
Pressure Sensor
Primary Valve Secondary Valve
MANUAL MODE
CAIS Switch ON
PMG
Aircraft in Flight NOTE Left cowl anti-ice system shown. Right cowl anti-ice system similar.
LEFT ENGINE Swirl Nozzle Location
Exhaust Slots
CAIS Valve Open CS1_CS3_3021_010
6th Stage Air
Figure 21: Cowl Anti-Ice System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-49
CS130_30_L3_CAIS_DetailedDescription.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the cowl anti-ice system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-50
CS130_30_L2_CAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
CAS MESSAGES Table 10: STATUS Messages
Table 8: CAUTION Messages MESSAGE
LOGIC
L ENG NACELLE OVHT
R ENG NACELLE OVHT
L COWL A/ICE FAIL
MESSAGE
Left burst cowl anti-ice duct or L HPC valve failed open or buffer air shutoff valve failed open. Right burst cowl anti-ice duct or R HPC valve failed open or buffer air shutoff valve failed open. Left cowl anti-ice system failed (valves closed).
R COWL A/ICE FAIL
Right cowl anti-ice system failed (valves closed).
L COWL A/ICE FAIL ON
Left cowl anti-ice system failed (valves open).
R COWL A/ICE FAIL ON
Right cowl anti-ice system failed (valves open).
COWL A/ICE ON
L or R COWL anti-ice manually selected ON while the OAT is above 15°C (59°F).
LOGIC
L COWL A/ICE ON
Left cowl anti-ice manually selected ON.
R COWL A/ICE ON
Right cowl anti-ice manually selected ON.
L-R COWL A/ICE ON
Left and right cowl anti-ice manually selected ON.
L COWL A/ICE OFF
Left cowl anti-ice manually selected OFF.
R COWL A/ICE OFF
Right cowl anti-ice manually selected OFF.
L-R COWL A/ICE OFF
Left and right cowl anti-ice manually selected OFF.
Table 9: ADVISORY Messages MESSAGE
LOGIC
L ENGINE FAULT
Loss of redundant or non-critical function for the left engine.
R ENGINE FAULT
Right of redundant or non-critical function for right engine.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-51
CS130_30_L2_CAIS_MonitoringTests.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
PRACTICAL ASPECTS COWL ANTI-ICE VALVES A manual override feature is provided on each valve to lock a failed valve in the open position for dispatch, which leaves the remaining valve for control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-52
CS130_30_L2_CAIS_PracticalAspects.fm
30 - Ice and Rain Protection 30-22 Cowl Anti-Ice System
Inlet
Solenoid Assembly
Outlet
Inlet
Manual Override
Solenoid Assembly
CS1_CS3_3021_004
MANUAL OVERRIDE
Outlet
Figure 22: Cowl Anti-Ice Valves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-53
CS130_30_L2_CAIS_PracticalAspects.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
34-10 PROBE HEAT SYSTEM A PROBE HEAT pushbutton annunciator (PBA) powers the ADSP and TAT probe heaters for 2 minutes on the ground provided the auxiliary power unit (APU) is running or external power is connected to the aircraft.
GENERAL DESCRIPTION The probe heat system prevents ice accumulation on the: •
Angle-of-attack (AOA) vanes
•
Air data system probes (ADSP)
•
Total air temperature (TAT) probes
WARNING
Each AOA vane has a vane heater and a case heater. For the left AOA vane, both heaters are powered by DC BUS 1. For the right AOA vane, both heaters are powered by DC BUS 2. These heaters are on whenever an engine is running or if the aircraft is in the air.
AS A SAFETY PRECAUTION, DO NOT TOUCH THE PROBES WITH BARE HANDS, AS THEY COULD BE HOT AND INFLICT BURNS.
The ADSP and TAT probe heaters operate in two modes: self-regulating and full. The self-regulating mode provides power to these heaters once the engines are running and the aircraft is on the ground under 55 kt. In this mode the probe regulates its own temperature. When the aircraft is in the air the heaters switch to full mode, which provides unregulated heating of the probes.
Table 11: Probe Heat Power Sources
The ADSP requires AC power for its primary and secondary power sources. Primary AC power is from AC BUS 1, AC BUS 2, and AC ESS BUS. Secondary power for ADSP 1 is AC BUS 2. For ADSP 2, secondary power is from AC BUS 1. The source of secondary power for ADSP 3 and 4 is from independent and dedicated static inverters which use power from the DC ESS BUS 3 to provide AC power. The primary heating power source for both TAT probes is the AC ESS BUS via their ADSPs. The left TAT probe heater is controlled by ADSP 3. The right TAT probe heater is controlled by ADSP 4. Secondary power is provided directly from the secondary power source from associated ADSPs.
PROBE
PRIMARY POWER
SECONDARY POWER
ADSP 1
AC BUS 1
AC BUS 2
ADSP 2
AC BUS 2
AC BUS 1
ADSP 3
AC ESS BUS
DC ESS BUS 3 via static inverter 1
ADSP 4
AC ESS BUS
DC ESS BUS 3 via static inverter 2
TAT L
ADSP 3
-----
TAT R
ADSP 4
-----
AOA L
DC BUS 1
-----
AOA R
DC BUS 2
----CS1_CS3_TB30.001
ADSP and TAT heater messages are displayed on the engine indication and crew alerting system (EICAS) and monitored by the onboard maintenance system (OMS). The AOA vane heaters are not monitored.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-54
CS130_30_L2_PHS_General.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
DC BUS 1 AC BUS 1
AC BUS 2
Primary AC Power
Secondary AC Power
Case
Vane
EICAS OMS ADSP 1
LEFT AOA VANE
PROBE HEAT AC ESS BUS
Primary AC Power
2 Minute Timer
GND ON
DC ESS BUS 3
Secondary AC Power
LEGEND ADSP AOA EEC TAT
Air Data System Probe Angle-of-Attack Electronic Engine Control Total Air Temperature
ADSP 3
NOTE Left side shown. Right side similar.
STATIC INVERTER 1
CS1_CS3_3030_003
LEFT TAT
Figure 23: Probe Heat System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-55
CS130_30_L2_PHS_General.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
COMPONENT LOCATION The probe heat system consists of the following: •
Angle-of-attack (AOA) vanes
•
Air data system probes (ADSPs)
•
Total air temperature (TAT) probes
•
Static inverters
ANGLE-OF-ATTACK VANE There are two angle-of-attack (AOA) vanes, one on each side of the forward fuselage. AIR DATA SYSTEM PROBE The are four air data system probes (ADSPs), two on each side of the forward fuselage. TOTAL AIR TEMPERATURE PROBE There are two total air temperatures (TATs) probes, one on each side of the forward fuselage. STATIC INVERTER There are two static inverters, one on each side of the forward equipment bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-56
CS130_30_L2_PHS_ComponentLocation.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
AOA Vane
TAT Probe
LEGEND ADSP AOA TAT
NOTE
Air Data System Probe Angle-of-Attack Total Air Temperature
Left side probes shown. Right side similar.
CS1_CS3_3030_005
Air Data System Probes
Figure 24: Probe Heat System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-57
CS130_30_L2_PHS_ComponentLocation.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
CONTROLS AND INDICATIONS A PROBE HEAT pushbutton annunciator (PBA) powers the ADSP and TAT probe heaters for 2 minutes on the ground provided the auxiliary power unit (APU) is running or external power is connected to the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-58
CS130_30_L2_PHS_ControlsIndications.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
OVERHEAD PANEL
AURAL WARN
PROBE HEAT
INHB
GND ON
L SIDE
OFF
OFF
R SIDE OFF
CS1_CS3_3030_007
OFF
WINDOW HEAT L WSHLD R
Figure 25: Probe Heat Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-59
CS130_30_L2_PHS_ControlsIndications.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
DETAILED DESCRIPTION AIR DATA SYSTEM PROBE Each angle-of-attack (AOA) vane has a vane heater and a case heater. For the left AOA vane, both heaters are powered by DC BUS 1. For the right AOA vane, both heaters are powered by DC BUS 2. The control and distribution cabinets (CDCs) provide power to the heaters. The heaters turn on when electrical power is available on the associated bus, and at least one engine is running or the aircraft is in the air. The air data system probes (ADSPs) use AC power for primary and secondary power sources. Primary AC power is from AC BUS 1, AC BUS 2, or AC ESS BUS. Secondary power for ADSP 1 is AC BUS 2. For ADSP 2, secondary power is from AC BUS 1. The source of secondary power for ADSP 3 and 4 is from static inverters which use 28 VDC power from the DC ESS BUS 3 to provide AC power. AC power from AC BUS 1 and AC BUS 2 is provided through solid-state power controllers (SSPCs). AC power from the AC ESS BUS is through thermal circuit breakers. The static inverters receive 28 VDC through thermal circuit breakers on electrical power control 3 (EPC 3).
When the PROBE HEAT pushbutton annunciator (PBA) is pressed, the white GND ON light illuminates and powers the probe heaters ON for 2 minutes and automatically turns off. The ADS PROBE HEAT GND ON info message is displayed on EICAS when the PROBE HEAT function is operating. The PROBE HEAT function is inhibited on ground unless at least one engine is running and AC power is feeding the aircraft. An ADSP PROBE HEAT FAIL caution message is displayed on EICAS if the associated probe heater is detected failed. The ADSP ISI PROBE HEAT caution message is displayed if both ADSP 3 and ADSP 4 heaters fail. The ADSPs provide data on ADSP and TAT heater operation to EICAS and OMS through the data concentrator unit module cabinets (DMCs).
WARNING
Power for the total air temperature (TAT) heaters is provided through the ADSP, from the AC ESS BUS.
ENSURE PROBE HEAT IS DEACTIVATED BEFORE JACKING THE AIRCRAFT.
The ADSP heaters and TAT probe heaters are switched off whenever the aircraft is on the ground and both engines are off. Otherwise these heaters operate in one of two modes: self-regulating or full mode.
PUTTING THE AIRCRAFT INTO THE AIR MODE AUTOMATICALLY ACTIVATES THE ADSP, TAT AND AOA PROBE HEATERS. ENSURE COVERS ARE REMOVED. HOT PROBES CAN CAUSE INJURY.
The self-regulating mode is enabled once the engines are running, the aircraft is on the ground, and airspeed is less than 55 kt. The probe heat is regulated to maintain a lower probe temperature to avoid overheating. The full mode is enabled when the airspeed is greater than 55 kt or the aircraft is in the air. The ADSP and TAT probe heaters operate in full mode if the controller integral heater fails.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-60
CS130_30_L3_PHS_DetailedDescription.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
CDC 3-6-4 L AOA Vane Heat DC BUS 1 SSPC 15A
LEFT AOA VANE Vane Heater
Logic: Combinational CDC 3-5-2 L AOA Case Heat DC BUS 1 SSPC 10A
Case Heater PROBE HEAT
Logic: Combinational
EICAS
GND ON
CB-G2 10
ADSP 3 ADSP 3A Heat
Heater Power 1
CB-G1 Left TAT Heat
5
LEGEND ADSP AOA CDC DMC TAT
Heater Power
Power
STATIC INVERTER 1
CB-A2 50
LEFT TAT PROBE
Left TAT
LCBP
DC ESS BUS 3
OMS
Inverter 1
28 VDC
115 VAC
Heater Power 2
EPC 3
NOTE • ADSP 3 shown, ADSP 1, 2, and 4 similar. • ADSP 1 primary heater power source is AC BUS 1, and the secondary heater power source is AC BUS 2. • ADSP 2 primary heater power source is AC BUS 2 and the secondary heater power source is AC BUS 1. • ADSP 4 primary heater power source is the AC ESS BUS and the secondary heater power source is DC ESS BUS 3 through static inverter 2. • AOA 2 heater power source is DC BUS 2. • Right TAT probe heater power source is ADSP 4.
Air Data System Probe Angle-of-Attack Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Total Air Temperature ARINC 429 Discrete
CS1_CS3_3031_001
AC ESS BUS
DMCs
Figure 26: Probe Heat Electrical Interface Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-61
CS130_30_L3_PHS_DetailedDescription.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages associated with the probe heat system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-62
CS130_30_L2_PHS_MonitoringTests.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
CAS MESSAGES Table 13: ADVISORY Messages
Table 12: CAUTION Messages MESSAGE
LOGIC
MESSAGE
ADS 1 PROBE HEAT FAIL
ADS 1 probe not heated per expectations.
ADS 2 PROBE HEAT FAIL
ADS 2 probe not heated per expectations.
ADS 3 PROBE HEAT FAIL
ADS 3 probe not heated per expectations.
ADS 4 PROBE HEAT FAIL
ADS 4 probe not heated per expectations.
ADS ISI FAIL ADS SAME SOURCE
LOGIC
ADS FAULT
Loss of redundant or non-critical function for the air data system.
ADS 1 DEGRADED
SSEC correction lost and based on default input value(s) for ADS 1 - includes loss of AOA offset.
Both ADS channels of ISI failed (channel 3 and 4).
ADS 2 DEGRADED
Combination of two displays using same source of air data (cover loss of channels 1-2, 1-3, 2-3, 1-4, 2-4).
SSEC correction lost and based on default input value(s) for ADS 2- includes loss of AOA offset.
ADS 3 DEGRADED
SSEC correction lost and based on default input value(s) for ADS 3- includes loss of AOA offset.
ADS 4 DEGRADED
SSEC correction lost and based on default input value(s) for ADS 4- includes loss of AOA offset.
ADS 1 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 2 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 3 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 1 FAIL
Loss of ADS 1 channel.
ADS 2 FAIL
Loss of ADS 2 channel.
ADS 4 SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS.
ADS 4 FAIL
Loss of ADS 3 channel.
ADS ISI SLIPCOMP FAIL
Loss of all sideslip compensation capability (cross and diagonal) in ADS 3 and 4.
ADS 3 FAIL
Loss of ADS 3 channel.
DUAL ADS FAIL
2 ADS sources are failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Table 14: STATUS Message MESSAGE ADS PROBE HEAT GND ON
TECHNICAL TRAINING MANUAL
For Training Purposes Only
LOGIC PROBE HEAT PBA selected and active.
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30 - Ice and Rain Protection 34-10 Probe Heat System
Page Intentionally Left Blank
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30 - Ice and Rain Protection 34-10 Probe Heat System
Table 15: INFO Messages Table 15: INFO Messages MESSAGE
MESSAGE
LOGIC
34 ADS FAULT - ADS SENSE LINE HEATER 1 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS SENSE LINE HEATER 2 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS SENSE LINE HEATER 3 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS SENSE LINE HEATER 4 INOP
Loss of sense line heater capability. Potential for long-term moisture buildup.
34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 1 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 2 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 3 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS HEATER Loss of heater control redundancy due to 4 REDUND LOSS either loss of one controller circuit, or loss of one power input. 34 ADS FAULT - ADS 1 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 2 heater or failure/loss of ASDSP 2 OSP electronics.
34 ADS FAULT - ADS 2 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 1 heater or failure/loss of ASDSP 1 OSP electronics.
34 ADS FAULT - ADS 3 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 4 heater or failure/loss of ASDSP 4 OSP electronics.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
34 ADS FAULT - ADS 4 PRIMARY SIDESLIP COMP INOP
Failure/loss of ADSP 3 heater or failure/loss of ASDSP 3 OSP electronics.
34 ADS FAULT - L TAT HEATER INOP
Fuselage TAT not able to be heated on left side. ADS 1 and 3 revert to onside engine TAT (loss of redundancy only).
34 ADS FAULT -R TAT HEATER INOP
Fuselage TAT not able to be heated on right side. ADS 2 and 4 revert to onside engine TAT (loss of redundancy only).
34 ADS FAULT - ADS 1 TAT ELEMENT INOP
ADS 1 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - ADS 2 TAT ELEMENT INOP
ADS 2 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - ADS 3 TAT ELEMENT INOP
ADS 3 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - ADS 4 TAT ELEMENT INOP
ADS 4 loss of on-side fuselage TAT measurement (automatic reversion to onside engine TAT - loss of redundancy only).
34 ADS FAULT - L AOA VANE INOP
ADS 1 or 3 is reporting a loss or internal failure (HW, SW, strapping changed since power-up) of the LEFT AOA VANE by setting SSM to failure warning or NCD.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_30_L2_PHS_MonitoringTests.fm
30 - Ice and Rain Protection 34-10 Probe Heat System
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30 - Ice and Rain Protection 34-10 Probe Heat System
Table 15: INFO Messages MESSAGE
LOGIC
34 ADS FAULT - R AOA VANE ADS 2 or 4 is reporting a loss or internal INOP failure (HW, SW, strapping changed since power-up) of the LEFT AOA VANE by setting SSM to failure warning or NCD. 34 ADS FAULT - L AOA VANE HEATER INOP
ADS 1 or 3 is reporting a heating failure of the LEFT AOA VANE. This INFO message is inhibited in case the AOA vane heater should not be heated per expectation.
34 ADS FAULT - R AOA VANE ADS 2 or 4 is reporting a heating failure of HEATER INOP the LEFT AOA VANE. This INFO message is inhibited in case the AOA vane heater should not be heated per expectation. 34 ADS FAULT - L AOA VANE DEGRADED
L AOA vane local AOA miscompares with ADSP 1 and ADSP 3 local AOA.
34 ADS FAULT - R AOA VANE L AOA vane local AOA miscompares with DEGRADED ADSP 2 and ADSP 4 local AOA. 34 ADS FAULT - L AOA CASE HEATER INOP
ADS 1 or 3 is reporting that the left AOA case heater current is lower than minimum for operational condition in flight. This INFO message is inhibited in case the AOA case heater should not be heated per expectation.
34 ADS FAULT - R AOA CASE ADS 2 or 4 is reporting that the left AOA HEATER INOP case heater current is lower than minimum for operational condition in flight. This INFO message is inhibited in case the AOA case heater should not be heated per expectation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_30_L2_PHS_MonitoringTests.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
30-41 WINDSHIELD AND SIDE WINDOW HEATING SYSTEM GENERAL DESCRIPTION The windshield and side window heating system (WWHS) applies electrical power to transparent films within the windshields and side windows. The windshields are heated to a higher temperature than the side windows to provide anti-icing and defogging. The side windows are heated to provide defogging only.
The WIPCs provide data on system operation for the engine indication and crew alerting system (EICAS) messages through the data concentrator unit module cabinets (DMCs).
All the windows are monitored for temperature by the associated windshield ice protection controller (WIPC) by using three temperature sensors embedded within each windshield and side window. During normal operation, one sensor provides temperature control and the other two sensors monitor the performance of the first one. In case of a sensor failure, the WIPC switches to one of the remaining two sensors for control and the last one is used as a monitor. Each window contains a transparent heating film, which is embedded between glass laminations. The heating film is configured into three zones. The WIPC supplies 3-phase 115 VAC to the heating film. Each zone is powered by single phase AC power. Electrical power can be turned OFF to each windshield and side window individually using the associated pushbutton annunciator (PBA) on the WINDOW HEAT panel located on the overhead panel. The left windshield WIPC heater power is supplied from the AC BUS 1 and the control power is supplied from DC BUS 1. The right windshield WIPC heater power is supplied from AC BUS 2 and the control power is supplied from DC BUS 2. The left side window WIPC heater power is from the AC ESS BUS and the control power is supplied from DC ESS BUS 3. The right side window WIPC heater power is supplied from AC BUS 2 and the control power is supplied from DC BUS 2. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-68
CS130_30_L2_WSWHS_General.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
LEFT SIDE WINDOW
LEFT WINDSHIELD
2
RIGHT WINDSHIELD
2
RIGHT SIDE WINDOW
2
2
1
1 1
1 Temperature Signal
LEGEND
DMC WIPC
Phase-A Phase-B Phase-C Data Concentrator Unit Module Cabinet Windshield Ice Protection Controller ARINC 429
WIPC 1
3-Phase AC Power
WIPC 3
L SIDE OFF
3X Temperature Sensors DC ESS BUS 3
2
Heating Films
WINDOW HEAT L WSHLD R OFF
OFF
DMC 1
NOTE 1
WIPC 2
CONTROL
AC ESS BUS HEATER
DC BUS 1
AC BUS 1
CONTROL
HEATER
WIPC 4
R SIDE OFF
DMC 2
EICAS OMS
DC BUS 2
AC BUS 2
DC BUS 2
AC BUS 2
CONTROL
HEATER
CONTROL
HEATER
CS1_CS3_3041_001
3-Phase AC Power
Figure 27: Windshield and Window Heat System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-69
CS130_30_L2_WSWHS_General.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
COMPONENT LOCATION The windshield and side window heating system (WWHS) consists of: •
Left and right heated windshield
•
Left and right heated side window
•
Windshield ice protection controllers (WIPCs)
WINDSHIELD There are two electrically-heated windshields located on each side of the forward section of the flight deck. SIDE WINDOW There are two electrically-heated side windows, one located on each side of the flight deck. WINDSHIELD ICE PROTECTION CONTROLLERS Four windshield and ice protection controllers (WIPCs) are installed in the forward equipment bay. Each window is assigned to a dedicated WIPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
FORWARD EQUIPMENT BAY
Side Window
WINDSHIELD ICE PROTECTION CONTROLLERS (4X)
CS1_CS3_3040_003
Windshield
Figure 28: Windshield, Side Window, and Windshield Ice Protection Controllers Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-71
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30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
CONTROLS AND INDICATIONS Each window has its own pushbutton annunciator (PBA). The PBAs are located on the WINDOW HEAT panel in the flight deck overhead panel. The PBAs command their own windshield ice protection controller through the data concentrator unit module cabinets (DMCs) to turn off the windshield/window heat. The OFF legend illuminates on the PBA to confirm that the corresponding window heat function is disabled. Cycling the PBA to OFF then ON again resets the WIPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
OVERHEAD
L SIDE OFF
WINDOW HEAT L WSHLD R OFF
OFF
R SIDE OFF
CS1_CS3_3041_004
WINDOW HEAT PANEL
Figure 29: Window Heat Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-73
CS130_30_L2_WSWHS_ControlsIndications.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
DETAILED DESCRIPTION The windshield and side window heating system (WWHS) applies electrical power to transparent films within the windows. The WWHS is normally on when the aircraft is powered. Window heating can be individually shut off through selection of the associated PBA located on the overhead panel. The controlled temperature of the windshield is set between 45.5°C (114°F) and 51°C (124°F). The side window temperature is set between 36.7° C (98°F) and 43.9°C (111°F). The left side window receives heater power from AC ESS BUS and control power from DC ESS BUS 3. As a result, under emergency conditions where normal AC power is lost, only the left side window can be heated. The WIPCs receive phase of flight data from the data concentrator unit module cabinet (DMC) to moderate power requirements for regulating window heat. For example, at initial power-up in cold soaked conditions, the WIPC regulates window temperature to a lower value to avoid damage to frozen windshields and windows due to thermal stresses. In the event of window heat failure, the WIPC reports through the DMCs, the appropriate WINDSHIELD HEAT FAIL or WINDOW HEAT FAIL caution message for display on EICAS. If the signal from the WINDOW HEAT PBAs is lost or ARINC 429 communications is lost with the DMCs, the WIPC turns on and a WHSLD FAIL ON, or a SIDE WDW HT FAIL ON advisory message is displayed for the affected window. The WIPCs periodically performs built-in tests (BITs) that are capable of detecting electrical power faults, temperature sensor faults, heating film faults, and internal WIPC faults. Test result data is communicated to the onboard maintenance system (OMS) through the DMCs.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-74
CS130_30_L3_WSWHS_DetailedDescription.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
LEFT SIDE WINDOW
LEGEND
DMC WIPC
Phase-A Phase-B Phase-C Data Concentrator Unit Module Cabinet Windshield Ice Protection Controller ARINC 429
2
2
RIGHT WINDSHIELD
2
RIGHT SIDE WINDOW
2
2
1
1
NOTE 1
LEFT WINDSHIELD
1
1
3X Temperature Sensors
Temperature Signal
3-Phase AC Power
Heating Films CB-F1 DC ESS BUS 3
3
L WINDOW CTRL
WIPC 3
WIPC 1
WIPC 2
L CBP
WIPC 4 CDC 4-5-11 R Window CTRL DC SSPC 3A BUS 2
AC ESS BUS
CDC 4-10-1, 2, 3 R Window HTR L SIDE OFF
CDC 3-4-17 L WSHLD CTRL DC SSPC 3A BUS 1 Logic: Always On EPC 1 AC BUS 1
25
L WSHLD HTR
WINDOW HEAT L WSHLD R OFF
OFF
DMC 1
R SIDE OFF
AC BUS 2
SSPC 7.5A
Logic: Always On
DMC 2
CDC 4-4-18 R WSHLD CTRL DC SSPC 3A BUS 2 Logic: Always On
EICAS OMS
R EPC 2 WSHLD HTR 25
AC BUS 2
CS1_CS3_3041_008
CB-H24 L WINDOW HTR 7.5
3-Phase AC Power
Logic: Always On
Figure 30: Windshield and Window Heat System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-75
CS130_30_L3_WSWHS_DetailedDescription.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the windshield and side window heating system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-76
CS130_30_L2_WSWHS_MonitoringTests.fm
30 - Ice and Rain Protection 30-41 Windshield and Side Window Heating System
CAS MESSAGES Table 16: CAUTION Messages MESSAGE
LOGIC
L SIDE WDW HEAT FAIL
Heater for left side window failed.
R SIDE WDW HEAT FAIL
Heater for right side window failed.
L WSHLD HEAT FAIL
Heater for left windshield window failed.
R WSHLD HEAT FAIL
Heater for right windshield window failed.
Table 17: ADVISORY Messages MESSAGE
LOGIC
L WSHLD HEAT FAIL ON
Heater for left windshield failed operative.
R WSHLD HEAT FAIL ON
Heater for right windshield failed operative.
L SIDE WDW HT FAIL ON
Heater for left side window failed operative.
R SIDE WDW HT FAIL ON
Heater for right side window failed operative.
Table 18: STATUS Messages MESSAGE
LOGIC
L SIDE WDW HEAT OFF
Heater for left side window selected OFF.
R SIDE WDW HEAT OFF
Heater for right side window selected OFF.
L WSHLD HEAT OFF
Heater for left windshield window selected OFF.
R WSHLD HEAT OFF
Heater for right windshield window selected OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-77
CS130_30_L2_WSWHS_MonitoringTests.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
30-42 WINDSHIELD WIPERS SYSTEM GENERAL DESCRIPTION Each windshield is provided with a windshield wiper with separate power sources, wiper motors, and associated switches, arms, and wiper blades. The left wiper motor is powered by AC BUS 1 and the right wiper motor is powered by AC BUS 2.
The wiper motors run continuous built-in tests (CBIT). In the event of failure the electronic controller sends failure information through the data concentrator unit module cabinets (DMCs) to the onboard maintenance system (OMS). There are no crew alerting system (CAS) messages for the windshield wiper system.
The wiper motors have electronic controllers and are electrically connected to each other for synchronization. This allows them to move together when both windshield wiper switches are selected to the same speed. The synchronization feature does not prevent the other wiper from operating if one wiper motor fails. The wiper controller incorporates a thermal protection that removes power to the motor if the motor overheats. After a cool down period, the controller reapplies power to the motor. When not in operation, the wipers move to the parked position, resting vertically at the center post of the windshield frame. If a wiper motor fails, it moves to the park position. The windshield wiper operation can be reset by setting the rotary switch to OFF and back to any of the other positions. The winshield wiper switch has the following selections: •
OFF: The onside wiper moves to the vertical parked position
•
INT: One wiping cycle every 5 seconds
•
SLOW: 80 cycles per minute
•
FAST: 120 cycles per minute
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-78
CS130_30_L2_WWS_General.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
Right Wiper Arm
Park Position
RIGHT WIPER MOTOR ELECTRONIC CONTROLLER CDC 4-9-4 R Wiper Motor AC SSPC 5A BUS 2 Logic: Always On
Left Wiper Arm
LEFT WIPER MOTOR sync
ELECTRONIC CONTROLLER
DMC 1 and DMC 2
OFF
LEGEND DMC OMS
Data Concentrator Unit Module Cabinet Onboard Maintenance System ARINC 429 Discrete
Logic: Always On
WIPER OFF
INT SLOW
OMS
FAST
INT SLOW FAST
CS1_CS3_3042_002
WIPER
CDC 3-9-6 L Wiper Motor AC SSPC 5A BUS 1
Figure 31: Windshield Wiper System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-79
CS130_30_L2_WWS_General.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
COMPONENT LOCATION The windshield wiper system consists of the following: •
Wiper motors
•
Wiper arms
•
Wiper blades
WIPER MOTOR There are two wiper motors. Each motor is located below its windshield. WIPER ARM There are two wiper arms. Each wiper arm is attached to its windshield wiper motor. WIPER BLADE There are two wiper blades. A wiper blade is attached to each wiper arm.
CAUTION To avoid damage to the windshield do not operate windshield wipers on dry glass.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-80
CS130_30_L2_WWS_ComponentLocation.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
Wiper Blade
Wiper Arm
CS1_CS3_3042_003
Wiper Motor
Figure 32: Windshield Wiper Motor, Wiper Arm, and Wiper Blade Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
30-81
CS130_30_L2_WWS_ComponentLocation.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
CONTROLS AND INDICATIONS The WIPER selector switches are located on the left and right wiper panel on the overhead panel of the flight deck. The switches allow selection of three speeds: intermittent (INT), SLOW, and FAST.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-82
CS130_30_L2_WWS_ControlsIndications.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
OVERHEAD PANEL
READING
OFF
BRT
READING
WIPER OFF
WIPER OFF
INT SLOW FAST
OFF
BRT
INT SLOW
CS1_CS3_3042_004
FAST
Figure 33: Windshield Wipers System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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30-83
CS130_30_L2_WWS_ControlsIndications.fm
30 - Ice and Rain Protection 30-42 Windshield Wipers System
Page Intentionally Left Blank
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30-84
CS130_30_ILB.fm
ATA 36 - Pneumatic
BD-500-1A10 BD-500-1A11
36 - Pneumatic
Table of Contents 36-10 Bleed Air System ..........................................................36-2 General Description .........................................................36-2 Left and/or Right Engine Bleed Air...............................36-2 Auxiliary Power Unit Bleed Air .....................................36-2 External Air Source ......................................................36-2 Component Location .......................................................36-4 Precooler Exhaust Door...............................................36-4 Bleed Temperature Sensor..........................................36-4 Bleed Monitoring Pressure Sensor ..............................36-4 Precooler......................................................................36-4 Fan Air Valve ...............................................................36-4 Intermediate Pressure Check Valve ............................36-4 High-Pressure Valve ....................................................36-4 Pressure-Regulating and Shutoff Valves .....................36-4 Cross-Bleed Valve .......................................................36-6 High-Pressure Ground Connection..............................36-6 Auxiliary Power Unit Check Valve................................36-6 APU Bleed Air Valve ....................................................36-8 Integrated Air System Controllers ..............................36-10 Detailed Component Information ....................................36-12 Integrated Air System Controller ...............................36-12 Electrical Interface .....................................................36-14 Controls and Indications .................................................36-16 Controls .....................................................................36-16 Indications .................................................................36-18 Detailed Description ......................................................36-20 Engine Bleed Air 4th and 8th Bleed Port Switching ..............................36-20 Pressure-Regulating Shutoff Valve Operation ...........36-22 Fan Air Valve Operation ............................................36-24 Precooler Exhaust Door.............................................36-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Bleed Air Source Selection ...................................... 36-28 Cross-Bleed Valve Operation.................................... 36-28 Monitoring and Tests .................................................... 36-30 CAS Messages ......................................................... 36-31 Practical Aspects ............................................................ 36-34 High-Pressure Valve, Pressure-Regulating Shutoff Valve, and Fan Air Valve Deactivation ......... 36-34 Cross-Bleed Valve Deactivation................................ 36-36 Precooler Exhaust Door Manual Opening and Closing.................................... 36-38 PCE Door Test .......................................................... 36-40 Onboard Maintenance System Tests ........................ 36-42 36-21 Bleed Air Leak and Overheat Detection System ....... 36-44 General Description ...................................................... 36-44 Bleed Air Leak and Overheat Detection System Operation .................................... 36-46 Component Location ..................................................... 36-48 Integrated Air System Controllers ............................. 36-48 Dual-Detection Loops................................................ 36-50 Manifolds................................................................... 36-50 Detailed Component Information ................................... 36-52 Detection-Loops ........................................................ 36-52 Manifolds................................................................... 36-52 Controls and Indications ................................................ 36-54 Detailed Description ...................................................... 36-56 Bleed Air Leak and Overheat Detection System Activation ..................................... 36-56 BALODS Leak Localization ...................................... 36-58 Monitoring and Tests ..................................................... 36-60 CAS Messages ......................................................... 36-61
TECHNICAL TRAINING MANUAL
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36-i
CS130_M5_36_PneumaticTOC.fm
36 - Pneumatic
List of Figures Figure 1: Bleed Air System Schematic .................................36-3 Figure 2: Engine Bleed Air System Components .................36-5 Figure 3: APU Check Valve, Cross-Bleed Valve, and High-Pressure Ground Connection Location.................................................................36-7 Figure 4: APU Bleed Air Valve .............................................36-9 Figure 5: Integrated Air System Controllers (IASCs) ..........36-11 Figure 6: Integrated Air System Controller ........................36-13 Figure 7: Electrical Interface...............................................36-15 Figure 8: AIR Panel ............................................................36-17 Figure 9: Bleed Air System Indications...............................36-19 Figure 10: Engine Bleed Port Switching ...............................36-21 Figure 11: PRSOV Operation ...............................................36-23 Figure 12: Fan Air Valve Operation ......................................36-25 Figure 13: Precooler Exhaust Door Operation .....................36-27 Figure 14: Bleed Air Source Selection and Cross-Bleed Valve Operation..............................36-29 Figure 15: High-Pressure Valve, Pressure-Regulating Shutoff Valve, and Fan Air Valve Deactivation ...........................36-35 Figure 16: Cross-Bleed Valve Deactivation..........................36-37 Figure 17: PCE Exhaust Door Manual Opening and Closing ..............................36-39 Figure 18: Precooler Exit Door Test .....................................36-41 Figure 19: Bleed Air System Initiated Built-In Test ...............36-43 Figure 20: BALODS Operation .............................................36-45 Figure 21: BALODS Zones...................................................36-47 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 22: Integrated Air System Controller......................... 36-49 Figure 23: Bleed Air Leak and Overheat Detection System Components .......................... 36-51 Figure 24: Manifold .............................................................. 36-53 Figure 25: Bleed Air Leak and Overheat Detection System Indications ............................. 36-55 Figure 26: Bleed Air Leak and Overheat Detection System Operation............................... 36-57 Figure 27: BALODS Leak Localization ............................... 36-59
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-ii
CS130_M5_36_PneumaticLOF.fm
36 - Pneumatic
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-iii
CS130_M5_36_PneumaticLOF.fm
PNEUMATIC - CHAPTER BREAKDOWN CHAPTER BREAKDOWN
Bleed Air System
1 Bleed Air Leak and Overheat Detection System
2
36 - Pneumatic 36-10 Bleed Air System
36-10 BLEED AIR SYSTEM GENERAL DESCRIPTION The aircraft bleed air system (BAS) receives hot bleed air from the engines, auxiliary power unit (APU), or an external air source. Pressurized air from these sources enters common ducting in the mid fuselage belly before being distributed to the user systems. A crossbleed valve (CBV) located in the manifold, isolates left and right side bleed air systems. Opening the CBV interconnects the two sides. The integrated air system controllers (IASCs) control and monitor the BAS. Each IASC is powered by two DC ESS BUSES. LEFT AND/OR RIGHT ENGINE BLEED AIR
valve (FAV) modulates the amount of engine fan air passing through the precooler in order to maintain the bleed air temperature limits. A bleed temperature sensor (BTS) provides the IASCs with temperature information to control the FAV.The fan air exhausts into the engine core compartment where it exhausts overboard. A precooler ex-haust (PCE) door opens automatically under certain high bleed air demand conditions to increase cool-ing airflow. Bleed air, downstream of the bleed air supply of both engines, is directed to a common high-pressure duct. A cross-bleed valve (CBV), mounted on the high-pressure duct, connects and isolates bleed air from both engines. The PRSOV and FAV are commanded closed by the IASC when the associated ENG FIRE PBA is pressed. Each valve has a thermal fuse that cuts its power if the temperature exceeds 250°C.
The left and right engines are the normal source of bleed air in flight. The BAS uses intermediate pressure bleed air from the 4th stage engine compressor, or high-pressure 8th stage bleed air supplied through the high-pressure valve (HPV). An intermediate pressure check valve (IPCKV) prevents 8th stage air from back flowing into the 4th stage port when the HPV is open.
AUXILIARY POWER UNIT BLEED AIR
Switching between 4th and 8th stage engine bleed air is based aircraft configuration and demand as calculated by the electronic engine control (EEC). The IASC uses this information to control the operation of the high-pressure valve (HPV). A bleed monitoring pressure sensor (BMPS) sends pressure information to the IASC for HPV position monitoring.
When the engines are started, the APU remains the primary bleed source to optimize engine performance during takeoff. When the APU performance decreases with altitude, the IASC automatically closes the APU BAV and switches the bleed source to the engines. When the APU bleed is selected off, the APU check valve prevents engine bleed air from reaching the APU.
Downstream of the BMPS, a pressure-regulating and shutoff valve (PRSOV) regulates bleed air pressure according to the bleed air demand requirements to a maximum of 59 psi. The torquemotor controlled PRSOV has a full closed microswitch for valve position reporting to the IASCs. The air conditioning pack inlet pressure sensors (PIPSs), (ATA 21), are used by the IASCs to regulate the pressure downstream of the PRSOV. A precooler, located downstream of the PRSOV control the engine bleed air temperature using engine fan air. A torquemotor controlled fan air Copyright © Bombardier Inc. CS130-21.07-06-00-121418
During normal ground operation, the IASC defaults the bleed air source to the APU. The APU bleed air valve (BAV) is controlled by the APU electronic control unit (ECU), (ATA 49). The IASC provides a load demand signal to the APU ECU. The ECU adjusts the APU load compressor output to meet the BAS demand.
EXTERNAL AIR SOURCE On the ground, the BAS can be supplied with pressurized air from an external air source, through the high-pressure ground connection (HPGC). The HPGC check valve closes when not in use to prevent bleed air system loss.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-2
CS130_36_L2_BAS_General.fm
36 - Pneumatic 36-10 Bleed Air System
L ENG
L EEC DC ESS 1
FIRE
DC ESS 2
IASC 1
DMC 1 and DMC 2
L PIPS 1 L PIPS 2
ENGINE AND APU FIRE PANEL
EICAS SYN OMS
T/M
Fan
FAV BMPS
T/M
PS
Precooler
To Right WAIS To Right Pack
To FTIS To Left Pack
BTS
M
From Right Engine Bleed Air System
4th Stage IPCKV 8th Stage
PRSOV
CROSS-BLEED VALVE
SOL
HPV
APU ECU
Overboard
APU CHECK VALVE
To Air Turbine Starter
BMPS BTS DMC ECU EEC FAV FTIS HPGC HPV
Fan Air Bleed Air Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Data Concentrator Unit Module Cabinet Electronic Control Unit Electronic Engine Control Fan Air Valve Fuel Tank Inerting System High Pressure Ground Connection High-Pressure Valve
IASC IPCKV PIPS PRSOV WAIS
Integrated Air System Controller Intermediate Pressure Check Value Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System ARINC 429
APU BLEED AIR VALVE
SOL
External Air
LEGEND
HPGC
APU
Motor
NOTE
SOL
Solenoid
Left Bleed Air System shown. Right Bleed Air System similar.
T/M
Torque Motor
M
CS1_CS3_3600_013
To Left WAIS
LEFT ENGINE
Figure 1: Bleed Air System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-3
CS130_36_L2_BAS_General.fm
36 - Pneumatic 36-10 Bleed Air System
COMPONENT LOCATION
BLEED MONITORING PRESSURE SENSOR
The main components of the bleed air system include:
The bleed monitoring pressure sensor (BMPS) is located in the pylon area. It is connected to the HP air duct downstream of the high-pressure shutoff valve, by a sense line.
•
Precooler exhaust (PCEs) doors
•
Bleed temperature sensors (BTSs)
•
Bleed monitoring pressure sensors (BMPSs)
•
Precoolers
•
Fan air valve (FAV)
•
Intermediate pressure check valve (IPCKV)
•
High-pressure valves (HPVs)
•
Pressure-regulating shutoff valves (PRSOVs)
•
Cross-bleed valve (CBV) (Refer to figure 3)
INTERMEDIATE PRESSURE CHECK VALVE
•
High-pressure ground connection (HPGC) (Refer to figure 3)
•
APU check valve (Refer to figure 3)
An intermediate pressure check valve (IPCKV) is located on the engine bleed air duct, at the 1 o’clock position.
•
APU bleed air valve (BAV) (Refer to figure 4)
HIGH-PRESSURE VALVE
•
Integrated air system controllers (IASCs) (Refer to figure 5)
The high-pressure valve (HPV) is located on the HP air duct, at the 2 o’clock position. It is also referred to as the high-pressure shutoff valve (HPSOV).
PRECOOLER The precooler, also known as the precooler exchanger is located on the engine pylon. The engine needs to be lowered in order to replace the precooler. FAN AIR VALVE The fan air valve (FAV) is located upstream of the precooler.
PRECOOLER EXHAUST DOOR The precooler exhaust (PCE) door is located on the engine right core cowl inner fixed structure, at the 3 o’clock position, aft of the oil tank access door.
PRESSURE-REGULATING AND SHUTOFF VALVES The pressure-regulating and shutoff valve (PRSOV) is located on the engine bleed air duct downstream of the HPV at the 1 o’clock position.
BLEED TEMPERATURE SENSOR The bleed temperature sensor is located on the high-pressure air duct, downstream of the precooler in the pylon area.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-4
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
Bleed Temperature Sensor
Bleed Monitoring Pressure Sensor (mounted on pylon)
Precooler Pylon
Fan Air Valve
Intermediate Pressure Check Valve
ENGINE RIGHT CORE COWL
High-Pressure Pressure-Regulating Valve Shutoff Valve
RIGHT ENGINE (OUTBOARD VIEW)
CS1_CS3_3615_002
Precooler Exhaust Door (access panel removed for clarity)
Figure 2: Engine Bleed Air System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-5
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
CROSS-BLEED VALVE The cross-bleed valve (CBV) is located in the mid forward section of the wing-to-body fairing (WTBF) area. HIGH-PRESSURE GROUND CONNECTION The high-pressure ground connection (HPGC) is located in the left forward section off the WTBF area. AUXILIARY POWER UNIT CHECK VALVE The auxiliary power unit (APU) check valve is located in the right aft section of the WTBF area.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-6
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
A B C
CROSS-BLEED VALVE
B HIGH-PRESSURE GROUND CONNECTION
C APU CHECK VALVE
CS1_CS3_3615_010
A
Figure 3: APU Check Valve, Cross-Bleed Valve, and High-Pressure Ground Connection Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-7
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
APU BLEED AIR VALVE The APU bleed air valve (BAV) is located on the APU.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-8
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
CS1_CS3_3615_015
AUXILIARY POWER UNIT (BOTTOM VIEW)
APU BLEED AIR VALVE
Figure 4: APU Bleed Air Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-9
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
INTEGRATED AIR SYSTEM CONTROLLERS Two integrated air system controllers (IASCs), IASC 1 and IASC 2 are located in the mid equipment bay forward rack.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-10
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
IASC 1 LEGEND IASC
Integrated Air System Controller
MID EQUIPMENT BAY FWD RACK
CS1_CS3_3615_017
IASC 2
Figure 5: Integrated Air System Controllers (IASCs) Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-11
CS130_36_L2_BAS_ComponentLocation.fm
36 - Pneumatic 36-10 Bleed Air System
DETAILED COMPONENT INFORMATION INTEGRATED AIR SYSTEM CONTROLLER Each IASC incorporates a channel A and a channel B. Channel A is the primary control channel, and channel B is the standby channel. On the ground the IASC 1 alternates with IASC 2 based on odd and even days. If a failure is detected on the primary control channel, the IASC switches to the healthy channel. For safety purposes, a third independent and dissimilar channel is integrated and referred to as the safety channel C. This allows system control independence from system failures annunciation. IASC channel functions are assigned accordingly: •
The left and right bleed pressure and temperature control is performed by the associated IASC primary control channel A or channel B
•
The left and right bleed pressure monitoring is performed by the associated IASC primary control channel A and channel B
•
The left and right bleed temperature monitoring is performed by the associated IASC channel C, and the opposite IASC channel C. Temperature monitoring is also performed by all channels for auto-shutdown purposes
•
The CBV control and monitoring is performed by the left IASC channel B, and the right IASC channel B
•
The APU bleed control is performed by the left and right IASC primary control channel A or channel B
An initiated built-in test (IBIT) of the bleed air system (BAS) can be performed from the onboard maintenance system (OMS) for each channel and controller.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-12
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
IASC 1
Channel A
Channel B
Safety Channel C
• Left engine bleed management control and monitoring
• Left engine bleed management control and monitoring
• Left and right engine bleed management monitoring
• Left engine bleed shutoff control and monitoring
• Left engine bleed shutoff control and monitoring
• Left and right temperature monitoring
• APU bleed control
• APU bleed control • Cross-bleed control and monitoring
EICAS Data Exchange
DMC
SYN OMS
IASC 2
Channel A
Channel B
Safety Channel C
• Right engine bleed management control and monitoring
• Right engine bleed management control and monitoring
• Left and right engine bleed management monitoring
• Right engine bleed shutoff control and monitoring
• Right engine bleed shutoff control and monitoring
• Left and right temperature monitoring
• APU bleed control
• APU bleed control CS1_CS3_3620_004
• Cross-bleed control and monitoring
LEGEND ARINC 429
Figure 6: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-13
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
ELECTRICAL INTERFACE Each IASC is electrically powered by two independent DC ESS BUSES: •
IASC 1 is powered by DC ESS BUS 1 and DC ESS BUS 2
•
IASC 2 is powered by DC ESS BUS 2 and DC ESS BUS 3
The IASCs power the high-pressure valve (HPV), fan air valve (FAV), cross-bleed valve (CBV), and the pressure-regulating shutoff valve (PRSOV). The control and distribution cabinet (CDC) provides the signal to enable PRSOV operation through the IASC. The CDC command is provided when the AIR panel BLEED PBA is selected and a minimum of 15 psi bleed air is available. The PRSOV is monitored by the pack inlet pressure sensors (PIPS). When the PRSOV is open, the PIPS provide pressure information to control the output of the PRSOV. The PIPS also provide monitoring for low-pressure and overpressure conditions. The PRSOV is monitored for the closed position by a microswitch. The HPV solenoid operated valve is controlled by the IASC. The opening and closing of the HPV is based on intermediate pressure temperature information calculcated by the electronic engine control (EEC) and the bleed monitoring pressure sensors (BMPSs). The BMPS, located downstream of the intermediate pressure port check valve is used to monitor the position of the HPV.
When a main engine start signal is received when the aircraft is on the ground: the following valves are closed for the duration of the start plus 30 seconds. •
Wing anti-ice valve (WAIV)
•
Left and right flow control valves (FCV)
•
Trim air pressure regulating valve (TAPRV)
•
Trim air shutoff valve (TASOV)
When the engine start is complete, the valves reopen. In flight, the 30 second delay is not used and the valves open immediately after the engine starts. The EEC provides the following information: •
Engine running
•
Main engine start
•
N2
•
Intermediate pressure port temperature
The EEC information is used to configure the bleed air system and control system operation.
The PIPS and BMPS are powered from the DC ESS BUSES. The bleed temperature sensors (BTS) provide temperature information to the IASC for monitoring and control of the FAV. The IASC provides the required pneumatic demand information to the APU electronic control unit (ECU), when the APU is supplying bleed air. The APU load compressor provides the required output through the APU bleed air valve (BAV). The APU BAV is powered by the APU ECU.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-14
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
EPC 2 DC ESS BUS 2
IASC 2 IASC 2A
15
CH 2A
R CBP DC ESS BUS 3 CDC 1-6-10 L ECS SNSR DC ESS SSPC 3A BUS 1
15
L BTS IASC 2B
CH 2B CH S
L BLEED TEMPERATURE SENSOR
IASC 1 L EEC L PACK INLET PRESSURE SENSOR 1
L PIPS 2
L PACK INLET PRESSURE SENSOR 2
L PIPS 2 APU Bleed Enable
L BLEED MONITORING PRESSURE SENSOR
L BMPS
L BTS
Logic: Always On
Main Engine Start N2 Engine Running Intermediate Pressure Port Temperature
APU ELECTRONIC CONTROL UNIT
APU BLEED AIR VALVE
EPC 2 3
L BLEED SNSR
Main Engine Start Signal
Close Signal To: • WAIV’s • FCV’s • TAPRV • TASOV LEGEND
CDC 1 L FAV Torquemotor L PRSOV CMD
L FAN AIR VALVE
L PRSOV Enable L PRSOV Torquemotor L PRSOV Closed
L PRSOV
EPC 1 CBP DC ESS BUS 1
15
IASC 1A
EPC 2 CBP NOTE
DC ESS BUS 2
15
IASC 1B
CH 1A
L HPV SOL CMD
CBV Open/Closed CMD CBV Open CH 1B CBV Closed CH S
HIGH-PRESSURE VALVE
CROSS-BLEED VALVE
Left Bleed Air System Shown. Right Bleed Air System Similar.
BMPS BTS CBV CDC CL/OP CMD EEC FAV FCV FC FO HPV IASC PRSOV SNSR SOL TAPRV TASOV WAIV
Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Close/Open Command Electronic Engine Control Fan Air Valve Flow Control Valve Full Close Full Open High-Pressure Valve Integrated Air System Controller Pressure-Regulating Shutoff Valve Sensor Solenoid Trim Air Pressure Regulating Valve Trim Air Shutoff Valve Wing-Anti-Ice Valve ARINC 429
CS1_CS3_3611_005
DC ESS BUS 2
Figure 7: Electrical Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-15
CS130_36_L3_BAS_DetailedComponentInfo.fm
36 - Pneumatic 36-10 Bleed Air System
CONTROLS AND INDICATIONS CONTROLS The bleed air system (BAS) is controlled from the AIR panel, located on the flight deck overhead panel. Bleed air valve (BAV) operation is controlled by pushbutton annunciators (PBAs) and a rotary selector. When the L BLEED PBA, the R BLEED PBA, and the APU BLEED PBA are in the ON position, the operation of the associated BAV is automatically managed by the IASCs. When a PBA is selected OFF, the valve is forced closed and the OFF legend illuminates.
The cross-bleed valve (CBV) is controlled by the rotary XBLEED switch. The CBV is controlled automatically by the integrated air system controller (IASC) when the XBLEED switch is in the AUTO position. The IASC opens the CBV when there is a single source of bleed air supply (one engine, APU or external air). The CBV is closed or open when the XBLEED switch is selected to MAN CLSD or MAN OPEN respectively. An ENG BLEED MISCONFIG caution message is displayed if the CBV is open with both engine bleeds on.
The amber L (R) BLEED PBA FAIL light illuminates for the following conditions: •
Overpressure
•
Low pressure
•
Loss of both BTS sensing elements
•
HPV failed open
•
PRSOV failed open
•
IASC 1 or IASC 2 failure
The corresponding L (R) BLEED FAIL caution message is displayed on the engine indication and crew alerting system (EICAS) page. An amber FAIL light illuminates in the APU BLEED PBA in the event of failure to supply APU bleed air or the loss of bleed leak detection capability. The APU BLEED FAIL caution message is displayed on EICAS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-16
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
RAM AIR OVERHEAD PANEL
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
OPEN
FAIL OFF R BLEED
APU BLEED
FAIL OFF
FAIL OFF
CS1_CS3_3620_002
BLEED AIR PANEL
Figure 8: AIR Panel Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-17
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
INDICATIONS The bleed air system (BAS) operation indication is displayed on the AIR synoptic page. Bleed air flow lines, actual duct temperature, valve position in normal operation, valve position in fail operation, and pressure display with symbols and color codes. The EXT AIR symbol is displayed on the AIR synoptic page when the air pressure is detected, and the engines and APU bleed air are not operating.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-18
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
FLOW LINES
STATUS
AIR IR
DOOR OR
ELECC
FLT CTR CTRLL
FUEL FUE EL
HYD YD
AVIONIC VIONIC
INFO
CB
Condition Normal flow. No flow. Abnormal flow (overheat, out of range pressure,leak).
VALVE POSITION NORMAL OPERATION Symbol
Condition
Symbol
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
15°CC
10 °C °
15 °CC
Open with flow.
Open with flow.
Open with no flow.
Open with no flow.
5 PSI
40 PSI
-PSI
Normal.
EXT AIR
R PACK
40
XBLEED
Left Pressure-Regulating Shutoff Valve
EXT AIR
Condition Displayed when pressure detected and engine/APU bleeds are off.
HI Right Bleed Air Pressure
PSI
DECLUTTERED SYMBOLOGY LOGIC
Abnormal.
HI
Left Bleed Air Pressure
Invalid.
Symbol
RAM AIR
Condition Closed.
Condition
15 °CC
TRIM AIR L PACK
Symbol
AFT
CARGO
VALVE POSITION FAIL OPERATION
Closed.
PRESSURE
FWD
CKPT 23 °C 22 °C
APU
Invalid. SYNOPTIC PAGE – AIR
40 PSI
Right Pressure-Regulating and Shutoff Valve
CS1_CS3_3620_005
Symbol
Figure 9: Bleed Air System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-19
CS130_36_L2_BAS_ControlsIndications.fm
36 - Pneumatic 36-10 Bleed Air System
DETAILED DESCRIPTION The active integrated air system controller (IASC) automatically performs engine bleed air 4th and 8th bleed port switching, and controls the operation of the pressure-regulating shutoff valve (PRSOV), when the associated BLEED PBA on the AIR panel is selected ON.
To switch back to the engine IP port, the IASC uses the IPP information from the EEC. Once the actual pressure at the IP port matches the IPP calculated pressure, the IASC closes the HPV. During, descent, and approach or steep approach, the HP/IP threshold is increased to prevent excessive bleed port switching.
ENGINE BLEED AIR 4TH AND 8TH BLEED PORT SWITCHING
Temperature
The high-pressure (HP)/intermediate-pressure (IP) switching selects the bleed air from the engine IP or HP port based on the system configuration (number of packs, WAIS on) and demand of the user systems. The IASC automatically selects the proper engine bleed port based on pressure requirements and temperature limitations.
The bleed port switching logic also considers the IP temperature (IPT) to maintain WAIS performance under single or dual bleed source operation. If the IPT is lower than the WAIS threshold of 230°C (446°F), the IASC commands the HPV open. The HPV closes if the IPT increases above 240°C (464°F),
Pressure requirements are based on the required bleed flow, user system configuration, and engine thrust settings as calculated by the electronic engine control (EEC).
To protect the bleed air system downstream of the engine ports, the IASC inhibits the IP/HP switching when the HP port temperature (HPT) is above the maximum temperature threshold of 490°C (914°F). The HPV closes if the HPT increases above
Temperature limitations are based on the maximum allowable system temperature, and the minimum temperature to provide satisfactory WAIS operation.
If the HPV is open, and the maximum temperature threshold of 500°C (932°F) is reached, the IASC closes the HPV.
The engine bleed port switching is done by electrical control of the HPV. Pressure The IASC opens the HPV based on the EEC intermediate pressure port (IPP) calculated pressure information. The EEC calculation is the minimum pressure necessary to provide the required total bleed airflow in the system. If the actual pressure at the IP port, as indicated by the BMPS goes below a specified threshold, the IASC opens the HPV and the HP port supplies the pressure. The threshold varies based on the system configuration and demand. The HPV downstream pressure is limited to 75 +/- 5 psi.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-20
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L ENG
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
FIRE
OPEN
R PACK XBLEED
LEGEND
L EEC
DMC 1 and DMC 2
FAIL OFF
ENGINE AND APU FIRE PANEL
EICAS
R BLEED APU BLEED
ACS APU ATS BMPS BTS CBV CDC CMD DMC FAV FTIS HPGC HPV IASC PIPS PRSOV WAIS
SYN
FAIL OFF
FAIL
OMS
OFF
AIR PANEL CDC 1
IASCs • WAIS • ACS
L PIPS 1 L PIPS 2
L PRSOV CMD
LEFT ENGINE T/M
Fan
M
Users • L ACS • L WAIS • FTIS
FAV
BMPS
PS
T/M
Precooler
BTS
SOL
Motor
SOL
Solenoid
T/M
Torque Motor
M From Right Engine Bleed Air System
4th Stage 8th Stage
Fan Air Bleed Air Air Conditioning System Auxiliary Power Unit Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Command Data Concentrator Unit Module Cabinet Fan Air Valve Fuel Tank Inerting System High-Pressure Ground Connection High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System AFDX ARINC 429
PRSOV
CBV
To ATS
APU BLEED AIR VALVE
SOL
NOTE
External Air
Left Bleed Air System shown. Right Bleed Air System similar.
HPGC
APU
CS1_CS3_3600_019
HPV
Figure 10: Engine Bleed Port Switching Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-21
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
PRESSURE-REGULATING SHUTOFF VALVE OPERATION
•
When the BLEED PBA is selected ON, the control and distribution cabinet (CDC) provides an enable signal to the IASC, allowing the PRSOV to open and supply engine bleed air.
The BTS detects an overtemperature (refer to Fan Air Valve Operation)
•
The bleed air leak and overheat detection system (BALODS) detects a bleed air leak in the downstream ducting
Once enabled, PRSOV operation is controlled by the IASC. The IASC uses the pressure signal supplied by the air conditioning pack inlet pressure sensors (PIPSs), to maximize air conditioning performance. The IASC modulates the torque motor of the PRSOV to provide bleed air at a pressure of 41.5 psi. Refer to ATA 21, Environmental Control for additional information on PIPS.
If the PRSOV fails in the open position, the amber FAIL light illuminates in the associated L (R) PBA.
During a cross-bleed main engine start, the IASC controls the operating engine PRSOV to maintain 45 psi for engine starting. The PIPSs on the operating engine side provides the pressure signal to the IASC. When the EEC signals that the engine is not running, the IASC closes the associated PRSOV and HPV. The PRSOV is commanded closed by any of the following conditions: •
BLEED PBA is selected OFF
•
During a start of the associated engine. The PRSOV remains closed for the duration of the start plus an additional 30 seconds during a ground start. In flight, the 30 second delay is removed.
•
No bleed air flow demand (no WAI demand or air conditioning demand), provided the RAM AIR PBA is not selected open. If the RAM AIR PBA is selected open, the PRSOV is prevented from closing in order to provide trim air to the trim air system. Refer to ATA 21, Air Conditioning for more information
•
The associated ENG FIRE PBA is selected. This signal is provided through the data concentrator unit module cabinets (DMCs)
•
Overpressure of 45 psi (if bleed demand) or 60 psi (if no bleed demand), for more than 15 seconds, as reported by the associated PIPS. This protects the bleed air ducting from high engine bleed pressure
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-22
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L ENG
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
FIRE
OPEN
R PACK XBLEED
LEGEND
L EEC
DMC 1 and DMC 2
FAIL OFF
ENGINE AND APU FIRE PANEL
EICAS
R BLEED APU BLEED
ACS APU ATS BMPS BTS CBV CDC CMD DMC FAV FTIS HPGC HPV IASC PIPS PRSOV WAIS
SYN
FAIL OFF
FAIL
OMS
OFF
AIR PANEL CDC 1
IASCs • WAIS • ACS
L PIPS 1 L PIPS 2
L PRSOV CMD
LEFT ENGINE T/M
Fan
M
Users • L ACS • L WAIS • FTIS
FAV
BMPS
PS
T/M
Precooler
BTS
SOL
Motor
SOL
Solenoid
T/M
Torque Motor
M From Right Engine Bleed Air System
4th Stage 8th Stage
Fan Air Bleed Air Air Conditioning System Auxiliary Power Unit Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Control and Distribution Cabinet Command Data Concentrator Unit Module Cabinet Fan Air Valve Fuel Tank Inerting System High-Pressure Ground Connection High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System AFDX ARINC 429
PRSOV
CBV
To ATS
APU BLEED AIR VALVE
SOL
NOTE
External Air
Left Bleed Air System shown. Right Bleed Air System similar.
HPGC
APU
CS1_CS3_3600_019
HPV
Figure 11: PRSOV Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-23
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
FAN AIR VALVE OPERATION The IASC uses feedback from the BTS to control the fan air valve (FAV) in order to maintain a BAS temperature range between 170°C (338°F) and 232°C (450°F). The temperature regulation setpoint is determined by the IASC based on the number of engines operating, the number of air conditioning packs operating, and wing anti-ice system (WAIS) operation. The nominal setpoint with air conditioning running is 200°C (392°F). With the L (R) BLEED PBA selected on and the associated engine operating, the IASC enables the FAV to open. When the engine bleed air supply is commanded closed automatically or the L (R) BLEED PBA is selected OFF, and the PRSOV is confirmed closed by the IASC, the FAV closes. Should the PRSOV fail in the open position, the FAV to stay open to allow fan air cooling to the precooler. The FAV closes simultaneously with the PRSOV when the associated ENG FIRE PBA is selected. If the BTS detects a temperature greater than 257°C (495°F) for more than 30 seconds, the L (R) BLEED OVHT caution message is displayed on EICAS and the FAIL light on the L (R) BLEED PBA illuminates. The IASC closes the associated PRSOV (and HPV). The associated BLEED PBA on the AIR panel is selected OFF to match the valve position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-24
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L PACK FAIL OFF
MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL OFF
LEGEND
R PACK XBLEED
FAIL OFF R BLEED
APU BLEED
FAIL OFF
L ENG
FAIL OFF
FIRE
AIR PANEL
EICAS DMCs ENGINE AND APU FIRE PANEL
OMS
CDC 1
ATS BMPS BTS CDC CMD DMC FAV HPV IASC PIPS PRSOV WAIS
Fan Air Bleed Air Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Control and Distribution Cabinet Command Data Concentrator Unit Module Cabinet Fan Air Valve High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Wing Anti-Ice System AFDX ARINC 429
L PRSOV CMD
IASCs
LEFT ENGINE T/M
FAV BMPS
PS
T/M
Precooler
To Bleed Air System
4th Stage 8th Stage
BTS
SOL
PRSOV
HPV
NOTE To ATS
Left Bleed Air System shown. Right Bleed Air System similar.
CS1_CS3_3600_020
Fan
Figure 12: Fan Air Valve Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-25
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
PRECOOLER EXHAUST DOOR The precooler exhaust (PCE) door is located on the right inner fixed structure (IFS) panel at the 3 o'clock position. The door actuator is controlled by channel A of the electronic engine control (EEC). Operation of the actuator unlatches the exhaust door which is pushed open by two pistons. The PCE door can be opened electrically by the EEC, or manually by rotating the latches in a counterclockwise direction using a 0.635 cm (0.250 in.) drive tool. The door can only be closed manually. The PCE door has square drive adaptors to operate the upper and lower pistons, actuator, and striker latch. When closing the PCE door, an actuator reset check hole is used to confirm the actuator is in the correct position before securing the upper and lower pistons. When the precooler demand is high, the door opens based on logic from the EEC. The door actuator receives a 28 VDC signal command from EEC when all the following conditions are true: •
Single bleed air operation (from IASC)
•
Wing anti-ice is on (from IASC)
•
Bleed air is being extracted from the engine high-pressure port (from IASC)
•
Precooler outlet bleed temperature from the IASCs is above threshold for more than 10 seconds as determined by the EEC. The EEC determines the temperature limit based on altitude
•
Aircraft in flight
The L (R) ENG PCE DOOR OPEN advisory message is displayed on EICAS when the precooler exhaust (PCE) door is open. The EEC monitors the door latch actuator commands and reports failures to the onboard maintenance system (OMS). The PCE door opening function is also tested from the OMS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-26
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
Upper Piston Square Drive
Precooler Exhaust Door
Actuator Square Drive
Actuator Reset Check Hole
Striker Latch
Lower Piston Square Drive
EICAS OMS LEFT EEC DMC 1 and DMC 2
DMC EEC HP IASC WOW
Air/Gnd
Data Concentrator Unit Module Cabinet Electronic Engine Control High-Pressure Integrated Air System Controller Weight-On-Wheels ARINC 429 Discrete
Ch B Piston
IASC 1
Latch Actuator
NOTE Left engine PCE Door shown, right engine PCE Door similar.
CS1_CS3_3610_001
LEGEND
Ch A
Figure 13: Precooler Exhaust Door Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-27
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
BLEED AIR SOURCE SELECTION
•
Aircraft altitude is less than 16,000 ft
The IASC provides the logic for switching between the available bleed air sources.
•
Ice is detected above 1,500 ft or 2 minutes after takeoff
When the engines and the APU are not running, or their associated bleed switches are off, the IASC assumes a no bleed state. Pneumatic air from an external air source can then be used to supply air to the BAS. When bleed pressure is detected with no engine or APU source selected, the EXT AIR symbol appears on the AIR synoptic page. The IASC selects the APU as the default bleed air source, when the APU electronic control unit (ECU) provides a ready-to-bleed signal, and all of the following conditions apply: •
The APU BLEED PBAs is in the on (normal) position
•
The WING ANTI-ICE switch is in the AUTO position and the system has not been activated.
•
Slats and flaps extended (not fully retracted)
•
Aircraft altitude <15,500 ft
In addition, whenever both engines are selected OFF (or not running), as signaled from the EECs, the IASCs switch to the APU as the default bleed air source. When the IASC selects APU as the bleed air source, it signals the APU ECU to open the APU bleed air valve (BAV) and the cross-bleed valve (CBV). During and after engine start, the APU remains the default bleed air source. It supplies the environmental control system (ECS) on the ground and during takeoff to minimize the bleed demand on the engines. The IASC automatically switches to engine bleed air, provided either or both engines are running and the associated BLEED PBA is in the AUTO position, and any of the following conditions apply: •
APU BLEED PBA selected OFF or the APU is not ready to bleed
•
WING ANTI-ICE switch selected to ON or in AUTO and active
•
Slats and flaps fully retracted
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
When the IASC switches to the engine bleed air source, the APU BAV closes through the APU ECU and both PRSOVs are opened. During descent, with the APU available, the bleed air system switches from the engines to the APU when the aircraft altitude is less than 15,500 ft, flaps and slats are extended, and wing anti-ice is not required. The ENG BLEED MISCONFIG caution message is displayed on EICAS if both engine BLEED PBAs are selected OFF, and a low bleed air flow is detected from the APU. This indicates that the APU is unable to support the bleed air users while the APU is the bleed source. CROSS-BLEED VALVE OPERATION The cross-bleed valve (CBV) is directly controlled by the IASCs channel B. With the XBLEED switch in the AUTO (normal) position, the IASC automatically controls the opening and closing for the CBV. The crew can force the IASC to command the CBV open or closed by selection of the MAN OPEN or MAN CLSD positions. When the CBV is manually operated, XBLEED MAN OPEN or XBLEED MAN CLSD EICAS status messages are displayed. In the AUTO position, the IASC commands the CBV open whenever a single bleed source is available (one PRSOV open, APU bleed air source or external air). The CBV is closed when both PRSOVs are open. The ENG BLEED MISCONFIG caution message is displayed on EICAS if the XBLEED valve is selected to MAN OPEN while both engine BLEED PBAs are in the ON position. If the IASC detects that the CBV has failed to move (failed in position) when commanded, the XBLEED FAIL caution message is displayed on EICAS. This situation can be detected whether the XBLEED switch is selected to AUTO or any of the MAN positions. Under these conditions the IASC latches the EICAS message until an initiated built-in test (IBIT) is performed.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-28
CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
L PACK
OFF
XBLEED AUTO
MAN CLSD
FAIL
MAN OPEN
OFF
L BLEED FAIL OFF
R BLEED
Left EEC and Right EEC
SFECU 1 and SFECU 2
ADSPs
FAIL
APU BLEED
EICAS
OFF
FAIL
DMC 1 and DMC 2
OFF
OMS
ANTI-ICE WING AUTO
AIR PANEL
5 -40 PSI PSI
XBLEED
OFF
ON
5 -40 PSI
ANTI-ICE PANEL L PIPS 1 L PIPS 2
L PIPS 1 L PIPS 2
IASC 1 and IASC 2
ACS ADSP ATS BMPS BTS CBV CMD DMC ECU EEC FAV FTIS HPGC HPV IASC PIPS PRSOV SFECU WAIS M
APU
SYNOPTIC PAGE – AIR
LEFT ENGINE Fan
Users • L ACS • L WAIS • FTIS
T/M
FAV BMPS
PS
T/M
PRECOOLER
BTS
Fan Air Bleed Air Air Conditioning System Air Data System Probe Air Turbine Starter Bleed Monitoring Pressure Sensor Bleed Temperature Sensor Cross-Bleed Valve Command Data Concentrator Unit Module Cabinet Electronic Control Unit Electronic Engine Control Fan Air Valve Fuel Tank Inerting System High-Pressure Ground Connection High-Pressure Valve Integrated Air System Controller Pack Inlet Pressure Sensor Pressure-Regulating Shutoff Valve Slat/Flap Electronic Control Unit Wing Anti-Ice System ARINC 429 Motor
SOL
Solenoid
T/M
Torque Motor
Users • R ACS • R WAIS BTS
M
T/M
PRECOOLER
T/M
PS
BMPS
FAV
4th Stage
RIGHT ENGINE Fan 4th Stage
SOL
PRSOV
CBV
PRSOV
SOL
8th Stage
8th Stage HPV
To ATS
To ATS SOL
External Air HPGC
APU ECU
APU
HPV
CS1_CS3_3600_022
FAIL
LEGEND
R PACK
Figure 14: Bleed Air Source Selection and Cross-Bleed Valve Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L3_BAS_DetailedDescription.fm
36 - Pneumatic 36-10 Bleed Air System
MONITORING AND TESTS The following page displays the crew alerting system (CAS) and INFO messages associated with the bleed air system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L2_BAS_MonitoringTests.fm
36 - Pneumatic 36-10 Bleed Air System
Table 3: STATUS Messages
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
MESSAGE
LOGIC
LOGIC
XBLEED MAN CLSD
Crossbleed valve manually selected and confirmed closed.
L BLEED FAIL
Left engine bleed failure or loss of left bleed leak detection or loss of both IASC 1 channels.
XBLEED MAN OPEN
Crossbleed valve manually selected and confirmed opened.
R BLEED FAIL
Right engine bleed failure or loss of right bleed leak detection or loss of both IASC 2 channels.
APU BLEED OFF
APU bleed selected OFF.
APU BLEED FAIL
The APU is not able to provide bleed air when requested or loss of bleed leak detection.
L BLEED OVHT
Left engine bleed overtemperature detected.
R BLEED OVHT
Right engine bleed overtemperature detected.
ENGINE BLEED MISCONFIG
XBLEED valve selected open and both engine bleed selected on or low flow on APU and left and right engine BLEED PBAs OFF.
XBLEED FAIL
XBLEED valve failed in position in AUTO or MAN mode.
Table 2: ADVISORY Messages MESSAGE AIR SYSTEM FAULT
Table 4: INFO Messages MESSAGE
LOGIC
36 L BLEED FAIL - L HPV FAIL CLSD
Left HPV detected failed closed by either left IASC channel.
36 L BLEED FAIL - L HPV FAIL OPEN
Left HPV detected failed open by either left IASC channel.
36 L BLEED FAIL - L PRESS REG SOV INOP
Left PRSOV causing a low or high-pressure.
36 L BLEED FAIL - L PRESS SOV FAIL OPEN
Left PRSOV detected failed open by either left IASC channel.
36 R BLEED FAIL - R HPV Right HPV detected failed closed by either right FAIL CLSD IASC channel.
LOGIC Loss of redundant or non-critical function for the air system (bleed or trim).
36 R BLEED FAIL - R HPV Right HPV detected failed open by either right FAIL OPEN IASC channel.
L ENG PCE DOOR OPEN L PCE door commanded opened (due to precooler overtemperature condition).
36 R BLEED FAIL - R PRESS REG SOV INOP
R ENG PCE DOOR OPEN R PCE door commanded opened (due to precooler overtemperature condition).
Right PRSOV detected failed open by either 36 R BLEED FAIL - R PRESS REG PRESS SOV right IASC channel. HPV FAIL OPEN
Table 3: STATUS Messages MESSAGE
LOGIC
L BLEED OFF
Left engine bleed selected OFF.
R BLEED OFF
Right engine bleed selected OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Right PRSOV causing a low or high-pressure.
36 L BLEED FAIL - L TEMP L BTS failure. SNSR INOP 36 R BLEED FAIL - R TEMP SNSR INOP
TECHNICAL TRAINING MANUAL
For Training Purposes Only
R BTS failure.
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36 - Pneumatic 36-10 Bleed Air System
Page Intentionally Left Blank
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36 - Pneumatic 36-10 Bleed Air System
Table 4: INFO Messages MESSAGE
LOGIC
36 AIR SYSTEM FAULT - L L BTS drift or single channel failure. BLEED TEMP SNSR REDUND LOSS 36 AIR SYSTEM FAULT R BLEED TEMP SNSR REDUND LOSS
R BTS drift or single channel failure.
36 AIR SYSTEM FAULT - L Left BMPS failure. BLEED MON PRESS SNSR INOP 36 AIR SYSTEM FAULT R BLEED MON PRESS SNSR INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Right BMPS failure.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L2_BAS_MonitoringTests.fm
36 - Pneumatic 36-10 Bleed Air System
PRACTICAL ASPECTS HIGH-PRESSURE VALVE, PRESSURE-REGULATING SHUTOFF VALVE, AND FAN AIR VALVE DEACTIVATION The high-pressure valve (HPV), pressure-regulating shutoff valve (PRSOV), and the fan air valve (FAV) can be secured closed for dispatch by rotating the manual override lever to the CLOSED position. The venting screw is used to secure the manual override lever in the closed position. Removing the vent screw from its stowed position vents the pneumatic actuator and prevents valve operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
Lanyard Venting Screw Insertion Point
Venting Screw HIGH-PRESSURE VALVE DEACTIVATED
CS1_CS3_3600_015
HIGH-PRESSURE VALVE
NOTE High-pressure valve shown, pressure-regulating shutoff valve and fan air valve similar.
Figure 15: High-Pressure Valve, Pressure-Regulating Shutoff Valve, and Fan Air Valve Deactivation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
CROSS-BLEED VALVE DEACTIVATION The cross-bleed valve (CBV) can be secured closed by rotating the manual override lever to the closed position and securing the lever with lockwire.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
Electrical Actuator
CS1_CS3_3600_023
Manual Override Lever
Figure 16: Cross-Bleed Valve Deactivation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
PRECOOLER EXHAUST DOOR MANUAL OPENING AND CLOSING The PCE door can be manually operated using a square drive tool. Open the PCE Door 1. Rotate the upper piston assembly square drive 90° counterclockwise. 2. Rotate the lower piston assembly square drive 90° counterclockwise.
Close the PCE Door 1. Rotate the upper piston assembly square drive 90° counterclockwise. 2. Rotate the lower piston assembly square drive 90° counterclockwise. 3. Rotate the striker latch square drive 60° counterclockwise.
3. Rotate the striker latch square drive 60° counterclockwise.
4. Rotate the actuator square drive 10.5° counterclockwise.
4. Manually pull the door to the open position. 5. Rotate the upper piston square drive 90° clockwise.
NOTE
6. Rotate the lower piston assembly square drive 90° clockwise. 7. Rotate the striker latch square drive 60° clockwise.
The square drive will return to initial position. 5. Manually push the pre-cooler exhaust door to the closed position. 6. Hold the pre-cooler exhaust door in the closed position and rotate the striker latch square drive 60° clockwise. 7. Insert a 1/8 in. allen key into the actuator reset check hole. Verify that the allen key can be inserted from 1.524 to 3.048 cm (0.6 to 1.2 in.). If not, repeat the previous steps. 8. Rotate the upper piston assembly square drive 90° clockwise. 9. Rotate the lower piston assembly square drive 90° clockwise.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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36 - Pneumatic 36-10 Bleed Air System
Upper Piston Square Drive Actuator Square Drive
Actuator Reset Check Hole
Striker Latch
CS1_CS3_3610_002
Lower Piston Square Drive
Figure 17: PCE Exhaust Door Manual Opening and Closing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
PCE DOOR TEST The PCE Door Test is used to detect any operational faults of the PCE system, including the electrical interface, solenoid, and door assembly.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
MAINT MENU EEC1-A-PCE DOOR TEST
1/4
*** WARNING *** PRE-COOLER EXT DOOR WILL OPEN DURING THIS TEST
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- FADEC IS POWERED
* NOTE * TEST WILL COMMENCE IMMEDIATELY, BUT THE DURATION IS UNTIL THE OPERATOR CONFIRMS DOOR OPERATION
PRESS TO START TEST
Write Maintenance Reports
Cancel
CS1_CS3_7321_025
PRESS THE START BUTTON BELOW TO INITIATE TEST
Figure 18: Precooler Exit Door Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-41
CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
ONBOARD MAINTENANCE SYSTEM TESTS On power-up, each channel of the IASCs performs a power-up built-in test (PBIT) to validate the input and system functions. During system operations each channel performs a continuous built-in test (CBIT) to validate input, active function operations, and proper active output operations. PBIT and CBIT monitor adjacent and opposite channel faults and failures, and store these in each channel’s memory for maintenance retrieval. Initiated built-in tests (IBIT) of the PRSOV, HPV, FAV, BTS and BMPS can be conducted on the OMS. Preconditions and instructions for the tests are provided on the test pages. The OMS also provides data pages for alarms status, valve position, and sensor status. Refer to ATA 45, Central Maintenance System for additional information.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-10 Bleed Air System
MAINT MENU
RETURN TO LRU/SYS OPS
IASC1 chA-IBIT BAS IBIT BAS 1/4 EQUIPMENT TESTED: PRSOV, HPV,FAV, BTS, BMPS PRECONDITIONS 123456-
MAKE MAKE MAKE MAKE MAKE MAKE
SURE SURE SURE SURE SURE SURE
WEIGHT ON WHEEL L BLEED SW IS SELECTED OFF R BLEED SW IS SELECTED OFF ALL ENGINE ARE OFF WING A/ICE SW IS SELECTED OFF APU BLEED SW IS SELECTED OFF TEST INHIBIT/INTERLOCK CONDITIONS
WEIGHT ON WHEEL LEFT BLEED SWITCH RIGHT BLEED SWITCH ENGINE STATUS WING A/ICE SWITCH APU BLEED SWITCH SWITCH
YYY YYY YYY YYYY YYY
Continue
Write Maintenance Reports
Cancel
CS1_CS3_3600_006
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Figure 19: Bleed Air System Initiated Built-In Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BAS_PracticalAspects.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
36-21 BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM GENERAL DESCRIPTION The purpose of the bleed air leak and overheat detection system (BALODS) is to protect the aircraft from hot bleed air leaks, and subsequent overheat damage. The BALODS is a continuous monitoring system that detects the location of hot air leaks or overheat, and automatically isolates or reconfigures the affected system. The BALODS uses dual overheat detection loops that monitor the high-pressure ducting and air conditioning bays.
After an overheat detection and automatic reconfiguration, the flight crew may manually reconfigure the affected system from the AIR control panel. The BALODS system is active when the integrated air system controllers (IASCs) are powered. The BALODS function of the IASC is controlled by channel B, which is powered by DC ESS BUS 2 for IASC 1 and DC ESS BUS 3 for IASC 2.
Upon detection of overheat in the affected zone, the integrated air system controllers (IASCs) automatically command one or more valves to close within the following systems: •
Bleed air system: Left pressure-regulating shutoff valve (PRSOV), right PRSOV and cross-bleed valve (CBV)
•
Wing anti-ice system: Left and right wing anti-ice valves (WAIVs)
•
Air conditioning system: Left flow control valve (FCV) and right FCV
•
Trim air system (a sub-system of air conditioning): Left trim air shutoff valve (TASOV), and the right trim air pressure-regulating valve (TAPRV)
•
Auxiliary power unit (APU) bleed air: APU bleed air valve
Indication of the overheat condition is displayed on the engine indication and crew alerting system (EICAS) messages and on the AIR synoptic page. Failure of bleed leak detection capability is indicated by the EICAS messages and FAIL lights on the AIR panel. The IASCs are also capable of identifying and recording the location of the leak for use in troubleshooting. This information is available through the onboard maintenance system (OMS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
RAM AIR
BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM Overheat Detection Loops
BLEED AIR SYSTEM
L PRSOV
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
R PRSOV
L PACK
CBV
FAIL
DC ESS BUS 2 IASC 1
OFF
APU BLEED AIR VALVE
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED FAIL
APU ECU
OFF
OPEN
FAIL OFF R BLEED FAIL
APU BLEED
OFF
FAIL OFF
AIR PANEL Overheat Detection Loops
EICAS IASC 1 and IASC 2
DMC 1 and DMC 2
SYN OMS L WAIV R WAIV
L FCV
DC ESS BUS 3 IASC 2
AIR CONDITIONING SYSTEM
R FCV L TASOV R TAPRV
LEGEND CBV DMC ECU FCV IASC PRSOV TAPRV TASOV WAIV
Cross-Bleed Valve Data Concentrator Unit Module Cabinet Electronic Control Unit Flow Control Valve Integrated Air System Controller Pressure-Regulating Shutoff Valve Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Wing Anti-Ice Valve ARINC 429
CS1_CS3_3620_010
WING ANTI-ICE SYSTEM
Figure 20: BALODS Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-45
CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM OPERATION
Upon detection of overheat in the affected zone, the IASCs automatically perform the following actions within 15 seconds:
The dual-detection loops are installed along the high-pressure air ducts and associated components, in the following eight zones:
•
•
•
•
•
•
Left bleed manifold zone: Located downstream of the left engine pylon firewall and along the left inboard wing leading edge to the cross-bleed valve, the fuel tank inerting system (FTIS) heat exchanger, the auxiliary power unit check valve, the high-pressure ground connection, the trim air shutoff valve (TASOV), and on to the left flow control valve (FCV) Right bleed manifold zone: Located downstream of the right engine pylon firewall and along the right inboard wing leading edge to the cross-bleed valve (CBV), the trim air pressure-regulating valve (TAPRV), and on to the right FCV Left wing anti-ice zone: Located downstream of the left wing anti-ice valve to the left wing telescopic duct. There are no detection loops in the slats Right wing anti-ice zone: Located downstream of the right wing antiice valve to the right wing telescopic duct. There are no detection loops in the slats. Left pack zone: Located downstream of the left FCV to the mix manifold and associated aft cabin ducting
•
Right pack zone: Located downstream of the right FCV to the mix manifold and associated forward cabin ducting
•
Trim air duct zone: Located downstream of the left TASOV and right TAPRV to the respective ducting of the flight deck, forward cabin, aft cabin, and forward cargo trim air valves. In contrast to the forward and aft cabin, where only the trim air ducting is monitored, the complete air conditioning ducting to the flight deck, is monitored
•
APU duct zone: Located downstream of the APU bleed air valve to the APU check valve
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
•
Left bleed manifold zone: -
If on engine bleed, close the left pressure-regulating shutoff valve (PRSOV) and close the cross-bleed valve (CBV)
-
If on APU bleed, the IASCs sets the APU bleed request, to the APU electronic control unit (ECU) to false (closing the APU bleed air valve) and command the CBV to close
Right bleed manifold zone: -
If on engine bleed, close the right PRSOV and close the CBV
-
If on APU bleed, close the CBV
•
Left wing anti-ice zone: Close both wing anti-ice valves (WAIVs)
•
Right wing anti-ice zone: Close both WAIVs
•
Left pack zone: Close the left flow control valve (FCV)
•
Right pack zone: Close the right FCV
•
Trim air duct zone: Close the left TASOV and the right TAPRV
•
APU duct zone: Set the APU bleed request to false. The ECU, upon receiving the signal, closes the APU bleed air valve
The valves are locked out until the appropriate bleed air system (BAS) pushbutton annunciator (PBA) is selected OFF.
TECHNICAL TRAINING MANUAL
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36-46
CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Flight Compartment Left Aft Cabin
Left Fwd Cabin
Right Aft Cabin
Right Fwd Cabin MIX MANIFOLD FORWARD CARGO HEATING
Pressure Bulkhead LPGC
Pressure Bulkhead ERAV
TAV
TAV
TAV
TAV
LEFT PACK
RIGHT PACK FCV
LEFT ENGINE
Pylon Firewall
FCV
TASOV
Pylon Firewall
TAPRV
FTIS HEAT EXCHANGER
Precooler
RIGHT ENGINE Precooler
Telescopic Duct
CBV
WAIV
WAIV
Telescopic Duct
Tailcone Firewall
HPGC LEGEND FCV FTIS HPGC LPGC TAPRV TASOV TAV WAIV
Flow Control Valve Fuel Tank Inerting System High-Pressure Ground Connection Low-Pressure Ground Connection Trim Air Pressure-Regulating Valve Trim Air Shutoff Valve Trim Air Valve Wing Anti-Ice Valve Check Valve
APU BLEED AIR VALVE
APU
CS1_CS3_3620_006
CBV ERAV
APU Duct Zone Left Bleed Manifold Zone Left Pack Zone Left Wing Anti-Ice Zone Right Bleed Manifold Zone Right Pack Zone Right Wing Anti-Ice Zone Trim Air Duct Zone Cross-Bleed Valve Emergency Ram Air Valve
Figure 21: BALODS Zones Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-47
CS130_36_L2_BALODS_General.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
COMPONENT LOCATION The principal components of the bleed air leak and overheat detection system (BALODS) are: •
Integrated air system controllers
•
Dual-detection loops (Refer to figure 23)
•
Manifolds (Refer to figure 23)
INTEGRATED AIR SYSTEM CONTROLLERS The integrated air system controllers (IASCs) are located in the mid equipment bay forward rack.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
IASC 2
IASC 1
CS1_CS3_3615_018
MID EQUIPMENT BAY FWD RACK
INTEGRATED AIR SYSTEM CONTROLLERS (2X)
Figure 22: Integrated Air System Controller Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-49
CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
DUAL-DETECTION LOOPS The dual-detection loops are located on high-pressure bleed air ducts and in areas of bleed air system components. MANIFOLDS The manifolds are located on the high-pressure air ducts and coupling covers.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Dual-Detection Loops
Manifold Coupling Cover
High-Pressure Air Duct
Coupling Cover Manifold
Manifold Wing Anti-Ice Valve
Manifold
Dual-Detection Loops
CS1_CS3_3615_012
LEFT WING ANTI-ICE DUCTING
NOTE Dual-detection loops and manifolds are similar in other areas of the aircraft high-pressure air ducts and components.
Figure 23: Bleed Air Leak and Overheat Detection System Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L2_BALODS_ComponentLocation.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
DETAILED COMPONENT INFORMATION DETECTION-LOOPS Dual-detection loops consist of parallel rows of sensing elements. The sensing elements are connected in series, either directly to the next element or using a wire to interconnect elements where there is little chance of exposure to high temperatures. The loops are identified as loop A and loop B. Both ends of each loop are connected to the integrated air system controller (IASC). The loops function independently to minimize false leak indications and improve dispatch reliability. The loops are powered by the integrated air system controller (IASC) with 15 VAC. When a loop is exposed to a hot air leak its resistance changes and is detected by the IASC. MANIFOLDS Dual-detection loops are installed on manifolds that are located on high-pressure air ducts and coupling covers. The purpose of the manifolds is to direct hot air leaks to the loops when a leak occurs. The loops are also routed over hot bleed air components in some zones to provide overheat damage protection. A hot bleed air leak in the duct assemblies is contained by the metal (stainless steel) shroud. Bleed air flows through the insulation air gaps and is routed to the manifold which directs the hot air onto the dual-detection loops, triggering an overheat condition. The manifolds on the coupling covers differ from those on the duct assemblies. A poppet valve (spring-loaded ball) is used to prevent nuisance warnings caused by the allowable natural leakage at the flange interfaces. This arrangement allows a minor buildup of pressure before the poppet valve opens releasing the hot air and heating the loops.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_36_L3_BALODS_DetailedComponentInfo.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Dual-Detection Loops Detection Loop
Dual-Detection Loops
Spring
Ball
Detection Loop Coupling Cover
High-Pressure Air Duct COUPLING COVER MANIFOLD
MANIFOLD Shroud
Air Gap
Coupling Cover Manifold
Coupling Cover
Band Clamp (ref)
LEGEND Blead Air Leak
CROSS-SECTION OF HIGH-PRESSURE AIR DUCT ASSEMBLY
CS1_CS3_3615_013
V-Band Coupling (ref)
Figure 24: Manifold Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-53
CS130_36_L3_BALODS_DetailedComponentInfo.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
CONTROLS AND INDICATIONS The detection of a bleed air leak or overheat is displayed on the AIR synoptic page. Green flow lines indicate the normal presence of bleed air. When a leak is detected, the affected zone flow line(s) turn amber. A LEAK caution message is displayed on EICAS for the affected zone. White flow lines indicate that the particular zone has been confirmed isolated by the selection of the corresponding PBA or switch to OFF.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-54
CS130_36_L2_BALODS_ControlsIndications.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
67$786
AIR
DOOR
ELEC
FLT CTRL
)8(/
HYD
AVIONIC
INFO
CB FLOW LINES
FWD
&.37 23 °C 22 °C
CARGO
15°C
Symbol
AFT
Condition
LO
22 °C 22 °C 22 °C -+ 2 22 °C -+ 2
Valve displayed in white and closed position.
10 °C
15 °C
After isolation flow line turns to white no flow.
15 °C RAM AIR
TRIM AIR
Overheat displayed by amber flow line.
53$&.
/3$&.
40
XBLEED
36,
CAB ALT $38
ó3 LDG ELEV
5500 8.0 560
RATE CREW OXY
100 2000
SYNOPTIC PAGE – AIR
CS1_CS3_3620_007
40 36,
Figure 25: Bleed Air Leak and Overheat Detection System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-55
CS130_36_L2_BALODS_ControlsIndications.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
DETAILED DESCRIPTION BLEED AIR LEAK AND OVERHEAT DETECTION SYSTEM ACTIVATION Loop A of each dual-detection loop is connected to IASC 1 channel B, and loop B of each dual-detection loop is connected to IASC 2 channel B. Loop A and loop B of each detection loop is supplied with 15 VAC, and the IASC monitors the resistance of the current passing through each loop. During an overheat condition, the resistance of loop A and loop B changes. At a predetermined temperature, the IASC signals an overheat condition and automatically closes the valves in the affected system. Normal Mode Operation Under normal operation, detection by both of the dual-detection loops trigger the IASC with an overheat condition. The IASC automatically shuts down the affected zone and the appropriate crew alerting system (CAS) message is displayed on EICAS.
the affected zone to eliminate the risk of a bleed air leak. A LEAK DET FAIL caution message is displayed on EICAS for the affected zone and the associated FAIL light (L BLEED FAIL, R BLEED FAIL, L PACK FAIL or R PACK FAIL light) illuminates on the AIR panel. The crew are required to confirm the shutdown of the affected zone (s) by selecting the appropriate switches OFF. Selecting the switch to OFF resets the valve closing latch in the IASC. BALOD Failure If channel B of both IASCs fail, the entire BALODS is lost and considered failed and all zones are shut down. The LEAK DET FAIL caution message is displayed on EICAS. In this event the L and R BLEED pushbutton annunciators (PBAs) are selected OFF. If in flight, the aircraft must fly at a lower altitude as air conditioning is no longer available. Built-in Tests
Degraded Mode Operation
The IASC performs a power-up built-in test (PBIT), continuous BIT (CBIT), and landing BIT (LBIT). Initiated BIT (IBIT) can also be performed.
The BALODS operates in a degraded mode if one of the IASCs channel B is inoperative, or if loop A or loop B of the dual-detection loop fails. Loop failure is determined if one loop is shorted for more that 5 minutes, or one loop provides high-resistance (open) for more than 60 seconds.
PBIT is performed on every cold start on the ground. However, this cannot be done when bleed air is extracted as the system performs a calibration test of the loops and recording the resistance of the detection loops. This resistance value is stored by the IASC for use in detecting the location of bleed air leaks.
The LEAK DET FAULT advisory message is displayed on EICAS, indicating the fault and lost redundancy.
CBIT continually checks the loops for shorts and open circuits (continuity). A failure is used to switch to the degraded mode. The failure is indicated by the LEAK DET FAULT advisory message.
In the degraded mode, overheat detection relies on the remaining single loop to trigger the overheat annunciation for the affected zone. Failure Mode Operation In the event of a dual-loop failure in a particular zone, detection capabilities for that zone are lost. The IASC automatically shuts down Copyright © Bombardier Inc. CS130-21.07-06-00-121418
While the PBIT does a complete system test (shorts, continuity, etc) and calibration, the L-BIT only checks for loop continuity and internal system 5 minutes after landing. A loop failure is indicated by the LEAK DET FAULT advisory message.
TECHNICAL TRAINING MANUAL
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36-56
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
RAM AIR TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK
R PACK XBLEED
FAIL
MAN CLSD
OFF
AUTO
MAN OPEN
L BLEED FAIL OFF
OPEN
EICAS
FAIL
DMC 1 and DMC 2
OFF R BLEED
SYN
FAIL
APU BLEED
OMS
OFF
FAIL OFF
AIR PANEL
Typical Zone Detection Loops
ANTI-ICE WING AUTO
OFF
15 VAC
ON
Loop A 15 VAC Loop B
CH A
CH B
CH B
IASC 1
LEGEND IASC
Integrated Air System Controller ARINC 429
CH A
IASC 2
DC ESS BUS 2
DC ESS BUS 3
CS1_CS3_3615_014
ANTI-ICE PANEL
Figure 26: Bleed Air Leak and Overheat Detection System Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-57
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
BALODS LEAK LOCALIZATION The BALODS performs leak localization using software embedded in the IASCs. Leak location accuracy is within 50.8 cm (20 in.). The operating principle is based on the linear electrical resistance of the sensing elements inner conductor. Upon leak detection, the increasing temperature decreases the resistance between the inner and outer conductor. Software in the IASC processes this change in resistance with total loop resistance without a leak, and uses a ratio of the two resistances to indicate a leak location. Calibration of the loop resistance is performed automatically on power-up except if power-up occurs during bleed air extraction. BALODS leak localization by maintenance personnel is obtained through the OMS and compared to a reference table in ATA 05-51-44 of the aircraft maintenance publication (AMP).
NOTE The bleed air extraction system MUST BE OFF during the LEAK LOCALIZATION test to prevent the radiant heat from the bleed extraction from affecting the resistance value of the loops.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-58
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
LEGEND Data Concentrator Unit Module Cabinet Integrated Air System Controller ARINC 429
IASC1 CH A - Sensor Status
Write Maintenance Reports ZONE
EICAS DMC 1 and DMC 2
SYN OMS
Typical Zone Dual-detection Loops 2
In
Loop A 15 VAC
CH B
CH A
Out 1 1 CH A
CH B
IASC 1
Loop B 15 VAC
In
2
Out IASC 2
NOTE 1
Loop A and Loop B shorted by exposure to hot air leak, and triggers a change in resistance.
2
Software in IASC uses a ratio of each loop total resistance without leak, and total resistance with leak, to determine the location of the leak.
RETURN TO FAULT MSGS
IASC1 CHB
IASC2 CHB
001 / 002
L L L L
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
R R R R
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
L L L L
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
R R R R
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
L L L L
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
%
R R R R
BLEED BLEED BLEED BLEED
LOC FAIL FAULT LEAK
0 NO NO NO
0 NO NO NO
% CS1_CS3_3615_016
DMC IASC
MAINT MENU
Figure 27: BALODS Leak Localization Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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36-59
CS130_36_L3_BALODS_DetailedDescription.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the bleed air leak and overheat detection system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
36-60
CS130_36_L2_BALODS_MonitoringTests.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
CAS MESSAGES Table 7: INFO Message
Table 5: CAUTION Messages MESSAGE
LOGIC
MESSAGE
LOGIC
36 LEAK DET FAULT - Loss of one detection loop in the wing anti-ice zone. LOOP REDUND LOSS
L BLEED LEAK
Left bleed leak detected.
R BLEED LEAK
Right bleed leak detected.
APU BLEED LEAK
APU bleed leak detected.
L PACK LEAK
Left pack leak detected.
R PACK LEAK
Right pack leak detected.
TRIM AIR LEAK
Trim air leak detected.
WING A/ICE LEAK
Wing anti-ice leak detected.
LEAK DET FAIL
Entire leak detection system failed and either L BLEED, R BLEED, or APU BLEED PBA selected ON.
Table 6: ADVISORY Message MESSAGE LEAK DET FAULT
LOGIC One leak detection loop failed. Refer to INFO messages.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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36-61
CS130_36_L2_BALODS_MonitoringTests.fm
36 - Pneumatic 36-21 Bleed Air Leak and Overheat Detection System
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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36-62
CS130_36_ILB.fm
ATA 47 - Liquid Nitrogen (Fuel Tank Inerting System)
BD-500-1A10 BD-500-1A11
47 - Liquid Nitrogen (Fuel Tank Inerting System)
Table of Contents 47-10 Fuel Tank Inerting System Generation.........................47-2
47-20 Fuel Tank Inerting System Distribution...................... 47-36
General Description .........................................................47-2 Component Location .......................................................47-4 Component Information ...................................................47-6 Heat Exchanger ...........................................................47-6 Air Separation Module ................................................47-8 Detailed Description ......................................................47-10
General Description ...................................................... 47-36 Component Location ..................................................... 47-38 Check Valves ............................................................ 47-38 Flame Arrestor .......................................................... 47-38 Flow Balancing Orifices............................................. 47-38 Maintenance Port Plugs ............................................ 47-38
47-00 Fuel Tank Inerting System Control and Indicating...........................................................47-12 General Description .......................................................47-12 Component Location .....................................................47-14 Inerting Control Unit ..................................................47-16 Component Information .................................................47-18 Inlet Isolation Valve....................................................47-18 Temperature Isolation Valve ......................................47-18 Dual-Flow Shutoff Valve ............................................47-18 Detailed Description ......................................................47-20 Fuel Tank Inerting System Interface ..........................47-20 Inerting Control Unit ...................................................47-22 Normal Operating Mode.............................................47-24 Health Monitoring Mode.............................................47-28 Inhibit Mode ...............................................................47-28 Latched Shutdown Mode ...........................................47-28 Maintenance Mode ....................................................47-28 Monitoring and Test .......................................................47-30 CAS Messages ..........................................................47-31 Practical Aspects ............................................................47-32 Inlet Isolation Valve Lockout (TBC)............................47-32 Onboard Maintenance System ..................................47-34
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-i
CS130_M5_47_NitrogenTOC.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System)
List of Figures Figure 1: Fuel Tank Inerting System Generation..................47-3 Figure 2: Fuel Tank Inerting System Generation Components .......................................47-5 Figure 3: Heat Exchanger Assembly ....................................47-7 Figure 4: Air Separation Module...........................................47-9 Figure 5: Fuel Tank Inerting System Generation Detailed Description ............................................47-11 Figure 6: Fuel Tank Inerting System Control and Indicating .........................................47-13 Figure 7: Fuel Tank Inerting System Control Components ...........................................47-15 Figure 8: Inerting Control Unit ............................................47-17 Figure 9: Fuel Tank Inerting System Control Components ...........................................47-19 Figure 10: Fuel Tank Inerting System Interface ...................47-21 Figure 11: Inerting Control Unit ............................................47-23 Figure 12: Fuel Tank Inerting System Start Sequence.........47-25 Figure 13: Flow Control, Overtemperature , and Overpressure Protection ............................47-27 Figure 14: Fuel Tank Inerting System Modes.......................47-29 Figure 15: Inlet Isolation Valve Lockout................................47-33 Figure 16: Onboard Maintenance System............................47-35 Figure 17: Fuel Tank Inerting System Distribution ...............47-37 Figure 18: Fuel Tank Inerting System Distribution Components .....................................47-39
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47-ii
CS130_M5_47_NitrogenLOF.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System)
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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47-iii
CS130_M5_47_NitrogenLOF.fm
LIQUID NITROGEN - CHAPTER BREAKDOWN (FUEL TANK INERTING SYSTEM)
LIQUID NITROGEN (FUEL TANK INERTING SYSTEM) CHAPTER BREAKDOWN
Fuel Tank Inerting System Generation
1 Fuel Tank Inerting System Distribution
2 Fuel Tank Inerting System Indication
3 Fuel Tank Inerting System Control
4
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
47-10 FUEL TANK INERTING SYSTEM GENERATION GENERAL DESCRIPTION The fuel tank inerting system (FTIS) includes an onboard inert gas generation system (OBIGGS). The OBIGGS uses engine bleed air, auxiliary power unit (APU) air, or air from a ground cart to supply nitrogen-enriched air (NEA) to the fuel tanks. The inert NEA floods the empty space above the fuel, thus reducing the flammability in the fuel tanks. The system is controlled automatically through the inerting control unit (ICU). The bleed air is routed to the ozone converter, through an inlet isolation valve (IIV), to reduce the ozone concentration. The air supplied by the ozone converter passes through the heat exchanger assembly where it is cooled by the ram air. Some bleed air bypasses the heat exchanger via the temperature control valve (TCV). The TCV modulates to maintain the heat exchanger outlet temperature within a range that provides optimal OBIGGS performance. A ground cooling fan provides cooling air flow through the heat exchanger when the aircraft is at low airspeeds. A ground cooling fan check valve closes when the fan is operating. The check valve ensures that air is drawn over the heat exchanger during fan operation. The cooled air is filtered by the air separation module (ASM) inlet filter. Water vapor from the supply airflow is removed through a coalescer within the filter. The water drains from the filter assembly into the aircraft belly through a drain orifice in the bottom of the filter bowl. The cooled and filtered bleed air is supplied to the air separation module (ASM), which removes the oxygen from the air and creates the NEA.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-2
CS130_47_L2_FTISG_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
DC BUS 1
Ram Air Inlet Temperature Control Valve
ICU
S/M
SOL
O2P
P
OZONE CONVERTER
SOL
AIR SEPARATION MODULE
Bleed Air IIV
SOL SOL
To Fuel Tanks
ASM INLET FILTER
HEAT EXCHANGER ASSEMBLY
Ground Cooling Fan
HEAT EXCHANGER COOLING CHECK VALVE
LEGEND
Ground Cooling Fan Run Signal to ICU Ram Air Exhaust
ICU IIV
Bleed Air Cooling Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Inerting Control Unit Inlet Isolation Valve
CS1_CS3_4710_008
DC BUS 1
Figure 1: Fuel Tank Inerting System Generation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-3
CS130_47_L2_FTISG_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
COMPONENT LOCATION The following components of the FTIS are installed in the wing-to-body fairing (WTBF): •
Heat exchanger
•
Ground cooling fan
•
Heat exchanger cooling check valve
•
Ozone converter
•
Air separation module inlet filter
•
Air separation module
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-4 CS130_47_L2_FTISG_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
Air Separation Module
Air Separation Module Inlet Filter Ozone Converter Heat Exchanger Cooling Check Valve Heat Exchanger
CS1_CS3_4710_010
Ground Cooling Fan
Figure 2: Fuel Tank Inerting System Generation Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-5 CS130_47_L2_FTISG_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
COMPONENT INFORMATION HEAT EXCHANGER The heat exchanger consists of a cross flow heat exchanger element, a temperature control valve (TCV), and associated bypass ducting. The heat exchanger regulates the temperature of the air entering the air separation module (ASM), by mixing hot bleed air with air cooled by the heat exchanger. The TCV is a butterfly valve driven by a stepper motor. The ICU commands the TCV based on temperature signals from the ASM inlet temperature sensor. Limit switches on the valve provide position indication to the ICU. If the stepper motor fails, the valve remains in the last position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-6 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
Cooled Air Out Ram Air Inlet
Bleed Air In Ram Air Outlet
CS1_CS3_4710_004
Temperature Control Valve
Figure 3: Heat Exchanger Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-7 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
AIR SEPARATION MODULE The air separation module (ASM) is constructed of thousands of hollow fiber membranes that run the length of the module. The ASM separates the nitrogen and oxygen components of the bleed air into individual gas streams, using a process known as selective permeation. Air contains 78% nitrogen, 21% oxygen, and 1% other gases. Each gas has a characteristic permeation rate that is a function of its ability to dissolve and diffuse through a membrane. This characteristic allows fast gases, such as oxygen, to be separated from slow gases, such as nitrogen. Bleed air enters the ASM inlet at a controlled pressure and temperature. The bleed air feeds through hollow fiber membranes. The oxygen permeates through the walls of the membrane and discharges through the oxygen-enriched air (OEA) outlet. The remaining nitrogen-enriched air (NEA) flows down the hollow fiber membranes, to the NEA outlet, and on to the fuel tanks.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-8 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
Nitrogen-Enriched Air Outlet
Oxygen-Enriched Air Outlet
Flow
Hollow Fiber
CS1_CS3_4710_009
Inlet
Figure 4: Air Separation Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-9 CS130_47_L2_FTISG_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
DETAILED DESCRIPTION If the amount of ram cooling air is insufficient, the ground cooling fan (GCF) provides the cooling air flow to the heat exchanger assembly. The ICU uses the aircraft mach number information from the data concentrator unit module cabinet (DMC) to determine when the GCF should be running. The GCF runs anytime the mach number is less than 0.2 Mach. The GCF is powered by a solid-state power controller (SSPC), therefore a turn ON request is sent to the DMCs to activate the SSPC. At startup, the ICU energizes the GCF before opening the inlet isolation valve (IIV). This ensures that the heat exchanger is capable of cooling the initial flow of bleed air to avoid nuisance fuel tank inerting system (FTIS) shutdowns. The GCF sends a run signal to the ICU to indicate that it is operating. The fan speed and overspeed protection is managed by the fan control software. The fan automatically shuts down when overspeed is detected and a fault signal is set. When the ICU has determined the bleed air supply is within the normal range, the IIV opens to allow bleed air flow into the FTIS. The temperature control valve (TCV) is modulated to maintain the proper air separation module (ASM) inlet temperature of 60°C (140°F). If the heat exchanger outlet temperature exceeds 60°C (185°F), the TCV is driven towards the closed position and is fully closed at 85°C (185°F).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-10 CS130_47_L3_FTISG_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-10 Fuel Tank Inerting System Generation
DMC 1 AND DMC 2
ICU
CDC 1-11-7 Fuel Inert CTLR DC SSPC 5A BUS 1
Ram Air Inlet Temperature Control Valve
Logic: Always On
S/M
OZONE CONVERTER
SOL
TIV
Logic: Single Input
SOL SOL
To Fuel Tanks DFSOV
ASM INLET FILTER
HEAT EXCHANGER ASSEMBLY CDC 1-10-1 Electric Ground Cooling Fan DC BUS 1 SSPC 15A
O2P AIR SEPARATION MODULE
Bleed Air IIV
OXYGEN SENSOR
PRESSURE SENSOR ASM INLET TEMPERATURE P SENSOR
Ground Cooling Fan
HEAT EXCHANGER COOLING CHECK VALVE
LEGEND
Ground Cooling Fan Run Signal to ICU Ram Air Exhaust
DFSOV ICU IIV TIV
Bleed Air Cooling Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Dual-Flow Shutoff Valve Inerting Control Unit Inlet Isolation Valve Temperature Isolation Valve ARINC 429
CS1_CS3_4710_005
SOL
HEAT EXCHANGER OUTLET TEMPERATURE SENSOR
Figure 5: Fuel Tank Inerting System Generation Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-11 CS130_47_L3_FTISG_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
47-00 FUEL TANK INERTING SYSTEM CONTROL AND INDICATING GENERAL DESCRIPTION The inerting control unit (ICU) receives power from DC BUS 1. The ICU powers all of the fuel tank inerting system (FTIS) electrical components, except the ground cooling fan (GCF), which also receives power from DC BUS 1. The ICU software is field loadable. The FTIS enters the normal operating mode when the aircraft is fully powered, bleed air is available, and the ICU power-up built-in test (BIT) is complete. Operation of the FTIS is fully automatic. The inlet isolation valve (IIV) is the primary ON/OFF control for the FTIS. The IIV isolates the FTIS from the pneumatic system. The valve moves to the closed position for a loss of power or pneumatic pressure. The temperature isolation valve (TIV), located at the inlet to the air separation module (ASM), is used to shut off air flow to the distribution system. This prevents hot air from contacting fuel or fuel vapor during an overtemperature condition. When an overtemperature is detected, the valve is commanded closed. The TIV is an electrically-controlled, pneumatically-operated gate valve. A microswitch provides position sensing to the ICU. The dual-flow shutoff valve (DFSOV) supplies the distribution system with nitrogen-enriched air (NEA) at a flow rate based on altitude and vertical speed. It also prevents fuel and fuel vapor migration to the FTIS when the system is not operating. The DFSOV outputs three flow rates. The low-flow solenoid is energized on the ground, and in climb and cruise. The medium-flow solenoid is used in approach and slow descent, and the high-flow rate is used in descent. The high-flow is created by energizing both solenoids. System monitoring is limited to crew alerting system (CAS) messages indicating a fault has occurred and the system is shut down. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Temperature and pressure sensors are used for control, system health monitoring, and provide overtemperature and overpressure shutdown protection. The heat exchanger outlet temperature sensor monitors the heat exchanger discharge temperature. The air separation module (ASM) inlet temperature sensor monitors the temperature of the bleed air entering the ASM. The oxygen sensor monitors the outlet of the ASM. The sampling port is located at the inlet to the dual-flow shutoff valve. If the oxygen level concentration is detected above the preset threshold, a crew alerting system (CAS) INFO message is displayed. System monitoring is limited to CAS messages that indicate a fault has occurred and the system is degraded or shut down. The FTIS has five operating modes: •
Normal operating mode - the FTIS provides NEA when bleed air is available
•
Health monitoring mode - the ICU performs a pressure and oxygen concentration data check of the NEA once during cruise
•
Inhibit mode - power is removed from all FTIS components. The system recovers to the normal operating mode when the inhibit condition no longer exists
•
Latched shutdown mode - an overtemperature, overpressure fault, or sensor fault is detected and the system shuts down
•
Maintenance mode - provides the capability to test and troubleshoot FTIS faults to the line replaceable unit (LRU) level
TECHNICAL TRAINING MANUAL
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47-12 CS130_47_L2_FTISCI_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
DC BUS 1
ICU
Ram Air Inlet Temperature Control Valve
S/M
SOL
O2P
P
OZONE CONVERTER
SOL
Bleed Air IIV
TIV HEAT EXCHANGER ASSEMBLY
Heat Exchanger Outlet Temperature Sensor
SOL SOL
AIR SEPARATION MODULE ASM Inlet Temperature Sensor
To Fuel Tanks DFSOV
Oxygen Sensor
Pressure Sensor
Ground Cooling Fan
HEAT EXCHANGER COOLING CHECK VALVE
LEGEND
Ground Cooling Fan Run Signal to ICU ASM DFSOV ICU IIV TIV
Ram Air Exhaust
Bleed Air Cooling Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Air Separation Module Dual-Flow Shutoff Valve Inerting Control Unit Inlet Isolation Valve Temperature Isolation Valve
CS1_CS3_4730_002
DC BUS 1
Figure 6: Fuel Tank Inerting System Control and Indicating Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-13 CS130_47_L2_FTISCI_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
COMPONENT LOCATION The following fuel tank inerting control and indicating system components are located in the wing-to-body fairing (WTBF): •
Inlet isolation valve
•
Temperature isolation valve
•
Dual-flow shutoff valve
•
Heat exchanger outlet temperature sensor
•
Pressure sensor
•
Air separation module (ASM) inlet temperature sensor
•
Oxygen sensor
•
Inerting control unit (refer to figure 8)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-14 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Dual-Flow Shutoff Valve
Oxygen Sensor Temperature Isolation Valve
Inlet Isolation Valve
Pressure Sensor
ASM Inlet Temperature Sensor
CS1_CS3_4730_010
Heat Exchanger Outlet Temperature Sensor
Figure 7: Fuel Tank Inerting System Control Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-15 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
INERTING CONTROL UNIT The inerting control unit (ICU) is located in the mid equipment bay.
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47-16 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
CS1_CS3_4730_008
MID EQUIPMENT BAY AFT RACK
INERTING CONTROL UNIT
Figure 8: Inerting Control Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-17 CS130_47_L2_FTISCI_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
COMPONENT INFORMATION INLET ISOLATION VALVE The inlet isolation valve (IIV) is a solenoid-controlled, pneumatically-operated butterfly valve. The valve has a position indicator that must not be used to move the valve. TEMPERATURE ISOLATION VALVE The temperature isolation valve (TIV) is an electrically-controlled, pneumatically-operated gate valve. A microswitch provides position sensing to the ICU. The TIV also has a position indicator. DUAL-FLOW SHUTOFF VALVE The DFSOV is a dual-poppet, 28 VDC electrically-controlled, pneumatically-operated valve. The poppets are spring-loaded closed when the solenoids are de-energized. Visual indicators show when the poppets are in the open or closed position. Microswitches provide position information to the ICU when the poppets are closed.
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47-18 CS130_47_L2_FTISCI_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Inlet
Position Indicator
Position Indicator
Valve Position Indicator
Outlet
INLET ISOLATION VALVE Position Indicator TEMPERATURE ISOLATION VALVE
CS1_CS3_4730_003
DUAL-FLOW SHUTOFF VALVE
Figure 9: Fuel Tank Inerting System Control Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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47-19 CS130_47_L2_FTISCI_ComponentInformation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
DETAILED DESCRIPTION FUEL TANK INERTING SYSTEM INTERFACE Communication with the aircraft systems is through the data concentrator unit module cabinet (DMC). The ICU communicates with DMC 1. If DMC 1 becomes unavailable, the ICU will switch to DMC 2. The ICU transmits system and component health information, and receives flight phase and conditions information that is used by the ICU to determine the necessary nitrogen-enriched air (NEA) flow. The ARINC 429 inputs received via the DMC are from: •
Air data system for environmental data and Mach number
•
Fuel quantity computer for refuel status
•
Integrated air system controller (IASC) for bleed pressure
•
Onboard maintenance system (OMS) for testing
The digital channel controls the ground for the IIV and the TIV. Each analog circuit controls the 28 VDC for its respective valve. Analog channel 1 controls the power to the IIV, and analog channel 2 controls power to the TIV. In normal operation, the digital channel will remove the grounds from the IIV, TIV, and DFSOV when an overtemperature is sensed. If the digital channel fails, and the heat exchanger outlet temperature sensor senses an overheat, analog channel 1 removes power from the IIV. If the ASM inlet temperature sensor detects an overheat, analog channel 2 removes power from the TIV. The temperature control valve (TCV) position is based on both the heat exchanger outlet temperature sensor, and the ASM inlet temperature sensor. The ASM inlet sensor is the primary driver of temperature, but the heat exchanger outlet sensor provides a safety function and commands the valve to full-closed at 85°C (185°F).
The ICU remains in inhibit mode while performing power-up BIT (PBIT). After completing a successful PBIT, the ICU begins energizing valves in a controlled startup sequence. As each valve is energized, a position sensor check is performed and other safety related parameters are checked. The ICU digital channel is the primary monitor of FTIS temperature. Two independent temperature monitoring analog backup channels are provided for shutdown of the FTIS in case of a digital channel failure.
The pressure sensor at the air separation module (ASM) inlet is used for health monitoring and detection of overpressure conditions that could damage system components. The oxygen sensor provides an indication of the oxygen concentration and absolute pressure of the nitrogen-enriched air (NEA) delivered by the inerting system. The output from the oxygen sensor is used to monitor the performance of the ASM.
The heat exchanger outlet temperature sensor monitors the heat exchanger discharge temperature. The sensor is a dual-element type, One element reports to analog channel 1 of the inerting control unit (ICU), and the other element reports to the digital channel. The ASM inlet temperature sensor monitors the air temperature entering the ASM. This temperature is used by the inerting control unit (ICU) for temperature control and health monitoring. It is a dual-element type. One element reports to analog channel 2 of the ICU, and the other element reports to the digital channel. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-20 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
CDC 1-10-1 Electrical Ground Cooling Fan DC BUS 1 SSPC 15A
GROUND COOLING FAN
INERTING CONTROL UNIT
Logic: Single Input
Digital Channel GCF Run Signal
CDC 1-11-7 Fuel Tank Inerting System DC BUS 1 SSPC 5A
Position
INLET ISOLATION VALVE
Ground
Logic: Always On
Analog Channel 1 28 VDC HEAT EXCHANGER OUTLET TEMPERATURE SENSOR
Temperature
Temperature 28 VDC Fuel Quantity Computer
TEMPERATURE CONTROL VALVE
Position Position
TEMPERATURE ISOLATION VALVE
Ground EICAS
28 VDC Pressure DMC 1 and 2
OMS
PRESSURE SENSOR
Analog Channel 2 28 VDC
ASM INLET TEMPERATURE SENSOR
Temperature Air Data
IASC
LEGEND ASM GCF IASC
Air Separation Module Ground Cooling Fan Integrated Air System Controller ARINC 429
Temperature O2 Signal 28 VDC 28 VDC Position
OXYGEN SENSOR DUAL-FLOW SHUTOFF VALVE
CS1_CS3_4700_006
HMU
Figure 10: Fuel Tank Inerting System Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-21 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
INERTING CONTROL UNIT The inerting control unit (ICU) has a digital channel that is the primary monitor of FTIS temperatures. Two independent temperature monitoring analog channels shut down the FTIS in case of failure of the digital channel. When the digital or analog monitor circuits detect a failure, the controller enters a latched shutdown mode. The ICU provides automatic control of the FTIS. Primary functions include: •
System startup
•
System control
•
Temperature control
•
Overtemperature protection
•
Overpressure protection
•
Health monitoring
•
System shutdown
•
System inhibits
Secondary functions are part of a maintenance mode, which includes activating and monitoring each FTIS component as commanded by the aircraft's maintenance computer.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-22 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
System Startup System Control Analog Channel 1
Temperature Control Overtemperature Protection Overpressure Protection
Digital Channel
Health Monitoring
Analog Channel 2
System Shutdown System Inhibits Maintenance Modes
CS1_CS3_4740_005
INERTING CONTROL UNIT
Figure 11: Inerting Control Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-23 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
NORMAL OPERATING MODE The normal operating mode consists of the following submodes: •
Starting sequence
•
Flow control
•
Overpressure protection
•
Temperature control and overtemperature protection
6. Check that the heat exchanger outlet temperature sensor does not indicate overtemperature. 7. Open the TIV. 8. Check that ASM inlet temperature sensor does not indicate overtemperature.
When initially energized with 28 VDC, the ICU remains in inhibit mode while performing power-up BIT (PBIT). After completing a successful PBIT, the FTIS ICU begins energizing valves in a controlled startup sequence. As each valve is energized, a position sensor check is performed and other safety related parameters are checked.
9. Open mid flow solenoid of the dual-flow shutoff valve (DFSOV) for 30 second purge and then open low flow solenoid of the DFSOV. 10. If bleed air is not available, enter the inhibit mode. 11. If bleed air is available, enter the operating mode and begin temperature and flow control.
If the amount of ram cooling air is insufficient, the ground cooling fan (GCF) provides the cooling air flow to the heat exchanger assembly. The GCF runs if the aircraft mach number is less than 0.2 M. When the ICU is satisfied with the initialization checks, and has determined the bleed air supply is within the normal range, the IIV opens to allow bleed air flow into the FTIS. The temperature control valve (TCV) is modulated to maintain the proper air separation module (ASM) inlet temperature and the FTIS is considered powered-up and operational.
NOTE The ICU uses a bleed pressure signal from the IASC to determine if there is enough bleed air available to open the IIV and start the fuel tank inerting process.The PBIT test runs even if bleed air is not available. Since the IIV, TIV, and DFSOV do not open if bleed air is not available, they are only monitored for the closed position. The bleed pressure signal is also used to determine if the IIV has failed.
Starting Sequence The ICU controls the system startup as follows: 1. Perform a power-up BIT (PBIT). 2. Begin temperature monitoring. 3. Command the GCF to run if the aircraft is on ground, and check for GCF feedback. 4. Cycle the TCV open and closed. Confirm valve closed. 5. Open the inlet isolation valve (IIV) and confirm valve position. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-24 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Power-up BIT
Temperature monitoring
Ground cooling fan Ground cooling fan run signal
Temperature control valve cycled open Temperature control valve closed
Inlet isolation valve open Inlet isolation valve position feedback INERTING CONTROL UNIT Heat exchanger outlet temperature sensor detects no overtemperature
Open temperature isolation valve Temperature isolation valve position feedback
Dual-flow shutoff valve mid-solenoid open
No bleed air FTIS enters inhibit mode
30 second purge Low solenoid open
Bleed air available FTIS enters normal operating mode
CS1_CS3_4730_009
Air separation module inlet temperature sensor detects no overtemperature
Figure 12: Fuel Tank Inerting System Start Sequence Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-25 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Flow Control
Temperature Control and Overtemperature Protection
The ICU energizes the appropriate dual-flow shutoff valve (DFSOV) solenoids based on aircraft altitude and vertical speed.
The ICU maintains the ASM inlet temperature at 60°C (140°F). The ICU uses ASM inlet temperature sensor and HE outlet temperature sensor information to modulate the temperature control valve (TCV).
When low-flow is required, only the low solenoid is energized. For mid-flow, the mid solenoid is energized. High-flow is achieved with both solenoids energized. During ground, climb, and cruise operations, the fuel tank inerting system (FTIS) operates in low-flow mode. During this time, the NEA reduces the fuel tank oxygen concentration and fills the fuel tank space, storing NEA to minimize the amount required to be generated during descent. As the aircraft climbs, the NEA will be venting overboard in order to maintain a pressure slightly above ambient in the tank.
An overtemperature sensed by the digital channel closes all of the valves. An overtemperature sensed by either of the analog channels results in its respective valve closing. The controller enters the latched shutdown mode when either the digital or analog monitor circuits detect an overtemperature. The temperature shutdown thresholds are as follows: Digital:
During descent, a vent inflow condition occurs to maintain ambient pressure. The FTIS enters the high-flow mode to ensure that there is adequate mixing of the NEA and vent air.
•
85°C (185°F) for 2 minutes
•
88°C (190°F)
Overpressure Protection
•
Analog: 90°C (194°F)
The ICU provides overpressure protection of the FTIS components and ducting. If an overpressure condition of 60 psi is detected at the inlet to the air separation module ASM, the FTIS is automatically shut down by closing the inlet isolation valve (IIV) and DFSOV. The temperature inlet valve (TIV) is held open for a short period of time to allow any trapped pressure to bleed off through the filter drain and ASM oxygen-enriched air (OEA) port and then closed. Overpressure protection of the fuel tank is provided by the fixed orifices in the DFSOV.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-26 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
HE OUTLET TEMPERATURE SENSOR B
ASM INLET TEMPERATURE SENSOR B
Digital Channel
DFSOV Ground Cruise Climb
Low-Flow Solenoid
ICU
Fast Descent IIV Mid-Flow Solenoid
Slow Descent
SOL
TIV
SOL
DFSOV
SOL SOL
ICU Analog Channel 1
Analog Channel 2
Flow Control HE OUTLET TEMPERATURE SENSOR A
ASM INLET TEMPERATURE SENSOR A
Temperature Control and Overtemperature Protection
LEGEND
ICU Overpressure Sensor
SOL
SOL
IIV
TIV
ASM INLET FILTER
PS SOL SOL
ASM DFSOV
Overpressure Protection
OEA Port Trapped Pressure
Trapped Pressure
CS1_CS3_4740_006
ASM DFSOV HE ICU IIV OEA TIV
Bleed Air Warm Air Oxygen-Enriched Air Nitrogen-Enriched Air Air Separation Module Dual-Flow Shutoff Valve Heat Exchanger Inerting Control Unit Inlet Isolation Valve Oxygen Enriched Air Temperature Isolation Valve
Figure 13: Flow Control, Overtemperature, and Overpressure Protection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-27 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
HEALTH MONITORING MODE Although oxygen sampling is done continuously, the FTIS enters the health monitoring mode one time per flight. This occurs during the operating mode when the aircraft is in the cruise flight phase and above 18,000 ft. When the health monitoring interval is initiated, the oxygen sensor will be turned on to measure the NEA oxygen concentration at the ASM outlet for a duration of 1 minute. A small gas sample, drawn from the outlet of the ASM, is fed to the oxygen sensor. The pressure and oxygen concentration is sent to the ICU, where it is used in conjunction with other FTIS and aircraft parameters to determine system health. An INFO message is displayed on the engine indication and crew alerting system (EICAS) when the oxygen concentration is out of limits over a period of several flights.
The ICU exits inhibit mode and starts normal FTIS operation when the conditions causing the inhibit mode are no longer present for at least 3 seconds. LATCHED SHUTDOWN MODE In the latched shutdown mode, all FTIS components de-energize with no automatic recovery. The ICU enters into latched shutdown mode if one of the following events occurs: •
Overtemperature
•
Overpressure
•
Loss of ARINC 429 BUS data
•
Any FTIS sensor defective, disconnected, or out of range
INHIBIT MODE
•
Valve failure
In the inhibit mode, electrical power is removed from the system. System operation is stopped, but returns to the normal operating mode after the inhibit conditions cease.
The FTIS is reset through the onboard maintenance system (OMS). The fault is stored in the ICU non-volatile memory (NVM). The FTIS reactivates after a successful built-in test (BIT) and power-up cycle.
The FTIS enters the inhibit mode upon receiving a request from aircraft systems. All FTIS components de-energize in the inhibit mode, but return to an operational state when a change occurs in the status of another system. The following conditions cause an inhibit:
MAINTENANCE MODE
•
Fuel quantity computer (FQC) in refuel mode
•
FQC in defuel mode
•
Main engine start
•
Inflight air-assisted engine start
•
One engine inoperative inflight only
•
Single environmental control system bleed with wing anti-ice on
•
Steep approach with wing anti-ice on
•
Loss of bleed air
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The maintenance mode is accessed through the OMS. To aid in troubleshooting, the FTIS faults occurring during normal operation are recorded in NVM within the ICU. These NVM faults include the values of related parameters at the time of the fault, and are downloaded through the OMS while the aircraft is on the ground. The initiated BIT provides the capability to test the electrical line replaceable units (LRUs).
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-28 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
Refueling Defueling
EICAS
Main Engine Start ICU
OMS
Inflight Engine Start One Engine Inoperative Inflight
HMU
18,000 ft
Inhibit
Single ECS Bleed with WAI On Steep Approach With WAI On Loss of Bleed Air
Health Monitoring Mode
Inhibit Mode
Overtemperature Overpressure Loss of A429 BUS
Latched Shutdown
OMS Reset
Start Up
OMS
Sensor Fault
Latched Shutdown Mode
Maintenance Mode
CS1_CS3_4740_007
Valve Fault
Figure 14: Fuel Tank Inerting System Modes Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-29 CS130_47_L3_FTISCI_DetailedDescription.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
MONITORING AND TEST The following page provides the crew alerting system (CAS) and INFO messages for the control and monitoring system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-30 CS130_47_L2_FTISCI_MonitoringTest.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
CAS MESSAGES
Table 2: INFO Messages Table 1: ADVISORY Message
MESSAGE
MESSAGE
LOGIC
FUEL INERTING FAULT
The fuel tank inerting system fault exists. See INFO message.
47 FUEL INERTING FAULT TEMP ISOL VLV INOP
The fuel tank inerting system is prevented from operating (latched shutdown) until successfully completing a power-up BIT and an IBIT is successfully completed because a fault has been detected with the performance of the FTIS temperature isolation valve.
47 FUEL INERTING FAULT DUAL FLOW SOV INOP
The fuel tank inerting system is prevented from operating (latched shutdown) until successfully completing a power-up BIT and an IBIT is successfully completed because a fault has been detected with the performance of the FTIS dual flow shutoff valve.
Table 2: INFO Messages MESSAGE
LOGIC
47 FUEL INERTING FAULT-FUEL The fuel tank inerting system is INERTING SHUTDOWN prevented from operating (latched shutdown).
LOGIC
47 FUEL INERTING FAULT-FUEL The FTIS oxygen sensor detects INERTING DEGRADED abnormally high oxygen percentage in the nitrogen-enriched air stream or oxygen sensor fault. 47 FUEL INERTING FAULT-FUEL A loss of redundancy in the ICU INERTING REDUND LOSS overtemperature analog backup circuit has occurred, but has not resulted in the FTIS becoming inoperative. 47 FUEL INERTING FAULT-FUEL For inability to confirm IIV or TIV position INERTING INOP or loss of FTIS communication or controller fault. 47 FUEL INERTING FAULT INLET ISOL VLV INOP
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The fuel tank inerting system is prevented from operating (latched shutdown) until successfully completing a power-up BIT and an IBIT is successfully completed because a fault has been detected with the performance of the FTIS inlet isolation valve.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-31 CS130_47_L2_FTISCI_MonitoringTest.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
PRACTICAL ASPECTS INLET ISOLATION VALVE LOCKOUT (TBC) The fuel tank inerting system (FTIS) can be deactivated by locking out the IIV. Access to the IIV is through a small panel in the wing-to-body fairing (WTBF). The valve is locked out by removing the deactivation plug from the stowed position and threading it into the deactivation port. The plug is attached to the valve by a lanyard. Additional procedures require the FUEL INERT CTLR circuit breaker to be locked out.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-32 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
OPE
N
Fuel Tank Inerting Access Panel
Deactivation Port
Deactivation Plug
NOTE Lockout FUEL INERT CTLR circuit breaker.
INLET ISOLATION VALVE
CS1_CS3_4730_007
INLET ISOLATION VALVE
Figure 15: Inlet Isolation Valve Lockout Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-33 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
ONBOARD MAINTENANCE SYSTEM The onboard maintenance system initiated built-in test (IBIT) can be carried out with or without bleed air. The IBIT without bleed air checks all electrically actuated LRUs, however, the IIV, TIV, and DFSOV are only checked for the closed position as these valves are pneumatically operated. To carry out a full test of the IIV, TIV, and DFSOV, the IBIT with bleed air is used.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-34 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-00 Fuel Tank Inerting System Control and Indicating
MAINT MENU
MAINT MENU
FTIS ICU-IBIT Without Bleed Air
FTIS ICU-IBIT With Bleed Air IBIT WITH BLEED AIR
IBIT WITHOUT BLEED AIR
1/4
1/4
*** WARNING *** PERSONNEL SHOULD STAY CLEAR OF EXHAUST VENT TO AVOID BREATHING UNHEALTHY FUMES PRECONDITIONS
PRECONDITIONS
1- MAKE SURE THAT AIRCRAFT IS ON-GROUND
1- MAKE SURE THAT AIRCRAFT IS ON-GROUND
2- MAKE SURE AIRCRAFT IS IN MAINTENANCE MODE
2- MAKE SURE AIRCRAFT IS IN MAINTENANCE MODE 3- SET BLEED AIR TO GREATER THAN 20 PSIG
TEST INHIBIT/INTERLOCK CONDITIONS
ON-GROUND STATE MAINTENANCE MODE STATE
TEST INHIBIT/INTERLOCK CONDITIONS
= =
ON-GROUND STATE = MAINTENANCE MODE STATE = BLEED AIR PRESSURE (PSIG) = SNNNNNNN
* INVALID OPERATION * CHECK PRECONDITIONS BEFORE PRESSING "CONTINUE"
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Cancel
Continue
Write Maintenance Reports
Cancel
CS1_CS3_4740_009
PRESSING "CONTINUE" AGAIN WITH INHIBITS CONDITION WILL ABORT THE TEST
Figure 16: Onboard Maintenance System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-35 CS130_47_L2_FTISCI_PracticalAspects.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
47-20 FUEL TANK INERTING SYSTEM DISTRIBUTION GENERAL DESCRIPTION The nitrogen-enriched air (NEA) distribution system connects the onboard inert gas generation system to the fuel tanks. The dual-flow shutoff valve controls the amount of inert air flow to the fuel tanks. During ground, climb, and cruise operations, the system operates in low-flow mode, generating NEA. During this time, the air space above the fuel is used for NEA storage. This minimizes the amount of NEA required to be generated during descent. During descent, the fuel tank inerting system (FTIS) operates in mid- to high-flow mode, based on the descent rate. When the dual-flow shutoff valve (DFSOV) closes, the valve provides protection against fuel entering the FTIS. The distribution line from the DFSOV to the fuel tank has two springloaded flapper check valves that prevent any fuel that may enter the distribution lines from flowing out of the tank to the FTIS generating system.
The distribution network consists of tubing arranged in symmetrical left and right paths from the entry point. Flow through each path is controlled by flow balancing orifices located in the dry bay. Each of the orifices in the wing and center tanks have unique part numbers and are not interchangeable. They are sized to ensure that NEA flow is evenly distributed to each tank under all flow conditions. To minimize the chance of fuel entering the lines through sloshing or gravitational effects, the tubing is located at high points. The NEA discharge outlets are located as far as possible from the fuel vent line openings, to ensure that the tank is adequately flooded before the NEA flows overboard through the vent scoop. Maintenance port plugs are used to inspect for the presence of fuel in the distribution lines. They also provide access points for conducting pressure tests of the distribution lines and check valves.
The line incorporates a flame arrestor, which prevents fire from reaching the fuel tanks. The flame arrestor consists of a metal canister that contains a grid, specifically designed to eliminate passage of any flame. A J-trap is located immediately downstream of the distribution check valve inside the fuel tank. The J-trap provides a collection point for any fuel that enters the distribution system. This configuration prevents fuel from coming in contact with the distribution check valve flapper, minimizing the chance of fuel entering the FTIS lines. Any fuel in the J-trap is purged from the line during system startup.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-36 CS130_47_L2_FTISD_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
SOL SOL
DFSOV UPSTREAM CHECK VALVE Flame Arrestor
DISTRIBUTION CHECK VALVE
J-Trap
Maintenance Ports
Dry Bay
Center Fuel Tank Left Fuel Tank
CS1_CS3_4720_004
Right Fuel Tank
LEGEND Flow Balancing Orifice – Center Tank Flow Balancing Orifice – Wing Tank Maintenance Port Plug
Figure 17: Fuel Tank Inerting System Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
47-37 CS130_47_L2_FTISD_General.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
COMPONENT LOCATION The following components are installed in the distribution system: •
Check valves
•
Flame arrestor
•
Flow balancing orifices
•
Maintenance port plugs
CHECK VALVES The distribution check valve and upstream check valves are located at the front spar outside of the fuel tank near RIB 0. FLAME ARRESTOR The flame arrestor is mounted between the distribution and upstream check valves. FLOW BALANCING ORIFICES There are two flow balancing orifices in each dry bay outboard of RIB 6. One orifice is for the center tank distribution line and one is for the wing tank distribution line. MAINTENANCE PORT PLUGS Two maintenance port plugs are located in the distribution line near the center tank front spar.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-38 CS130_47_L2_FTISD_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
RIB 6L
RIB 0
Center Fuel Tank Front Spar
RIB 6L
To Right Wing
RIB 0 Flame Arrestor
Upstream Check Valve
Center Tank Flow Balancing Orifice
Left Wing Tank Flow Balancing Orifice (right wing tank similar)
Distribution Check Valve
Maintenance Port Plug
J-Trap (ref.)
Maintenance Port Plug
CS1_CS3_4720_005
To Left Wing
Figure 18: Fuel Tank Inerting System Distribution Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-39 CS130_47_L2_FTISD_ComponentLocation.fm
47 - Liquid Nitrogen (Fuel Tank Inerting System) 47-20 Fuel Tank Inerting System Distribution
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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47-1 CS130_47_ILB.fm
ATA 49 - Airborne Auxiliary Power
BD-500-1A10 BD-500-1A11
49 - Airborne Auxiliary Power
Table of Contents 49-00 APU Construction .........................................................49-2 General Description .........................................................49-2 Left Side View ..............................................................49-2 Right Side View............................................................49-4 Accessory Gearbox Section.........................................49-6 49-10 APU Installation ............................................................49-8 General Description ..........................................................49-8 Component Location .......................................................49-8 Component Information .................................................49-10 Access Doors.............................................................49-10 Access Door Struts ....................................................49-12 APU Drains ...............................................................49-14 Detailed Component Information ....................................49-16 Mounts and Struts......................................................49-16 Eductor ......................................................................49-18 Exhaust .....................................................................49-20 APU Drains Schematic .............................................49-22 49-12 APU Air Intake ............................................................49-24 General Description .......................................................49-24 Air Inlet.......................................................................49-24 APU Air Inlet Door......................................................49-24 APU Air Inlet Door Operation.....................................49-26 Component Location .....................................................49-28 Air Inlet Door ..............................................................49-28 Air Inlet Door Actuator................................................49-28 Air Inlet Door Position Sensor....................................49-28 Controls and Indications .................................................49-30 Detailed Description ......................................................49-32 Monitoring and Tests .....................................................49-34 CAS Messages ..........................................................49-35 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Practical Aspects ........................................................... 49-36 APU Inlet Door Mechanical Rigging.......................... 49-36 OMS APU Inlet Door Rigging.................................... 49-38 APU Door Lockout .................................................... 49-40 Onboard Maintenance System APU Door Inhibit ....................................................... 49-42 49-90 APU Oil System ......................................................... 49-44 General Description ...................................................... 49-44 Component Location .................................................... 49-46 Lubricating Module.................................................... 49-46 Oil Cooler ................................................................. 49-46 Thermostatic/Pressure-Relief Valve.......................... 49-46 Low Oil Pressure Switch ........................................... 49-46 Oil Level Sensor........................................................ 49-46 Gearbox .................................................................... 49-48 Magnetic Chip Collector ............................................ 49-48 Oil Temperature Sensor............................................ 49-48 Air/Oil Separator........................................................ 49-48 Detailed Component Information ................................... 49-50 Lubricating Module.................................................... 49-50 Controls and Indications ................................................ 49-52 Detailed Description ...................................................... 49-54 Monitoring and Tests .................................................... 49-56 CAS Messages ......................................................... 49-57 Practical Aspects ........................................................... 49-58 Oil Servicing .............................................................. 49-58 49-30 APU Fuel System ...................................................... 49-60 General Description ....................................................... 49-60 Component Location ..................................................... 49-62 Fuel Control Unit ....................................................... 49-62
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-i
CS130_M5_49_APUTOC.fm
49 - Airborne Auxiliary Power
Flow Divider Assembly...............................................49-62 Fuel Manifolds and Fuel Nozzles...............................49-62 Plenum Drain Orifice..................................................49-62 Detailed Component Information ....................................49-64 Fuel Nozzles ..............................................................49-64 Detailed Description ......................................................49-66 Fuel Control Unit ........................................................49-66 Fuel Flow Divider and Fuel Flow Divider Solenoid ........................................................49-68 Controls and Indications .................................................49-70 49-40 APU Start and Ignition Systems .................................49-72 General Description .......................................................49-72 Component Location .....................................................49-74 Starter Motor and Clutch Assembly ...........................49-74 Ignition Unit ................................................................49-74 Igniter Plug and Igniter Lead......................................49-74 Detailed Description ......................................................49-76 APU Start Power........................................................49-76 TRU Start ...................................................................49-76 Battery Start ...............................................................49-76 49-50 APU Air Systems ........................................................49-78 General Description .......................................................49-78 Component Location ......................................................49-80 Surge Control Valve...................................................49-80 P2 Sensor ..................................................................49-80 Total Pressure Sensor ...............................................49-80 Differential Pressure Sensor ......................................49-80 Inlet Temperature Sensor ..........................................49-80 Inlet Guide Vane Actuator..........................................49-80 Bleed Air Valve ..........................................................49-80 Detailed Description .......................................................49-82 APU Air Interface .......................................................49-82 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
APU Air Control......................................................... 49-84 Monitoring and Test ....................................................... 49-86 CAS Messages ......................................................... 49-87 49-60 APU Control .............................................................. 49-88 General Description ...................................................... 49-88 APU ECU Control Functions ..................................... 49-90 APU Start .................................................................. 49-92 APU Shutdown.......................................................... 49-94 Component Location ..................................................... 49-96 Exhaust Gas Temperature Thermocouple ................ 49-96 Speed Sensor ........................................................... 49-96 Data Memory Module................................................ 49-96 Electronic Control Unit .............................................. 49-98 Detailed Component Information ................................. 49-100 Data Memory Module.............................................. 49-100 Detailed Description .................................................... 49-102 ECU Power ............................................................. 49-102 ECU Interface.......................................................... 49-104 APU Protective Shutdowns ..................................... 49-106 Monitoring and Tests .................................................. 49-108 CAS Messages ....................................................... 49-109 Practical Aspects ......................................................... 49-110 Onboard Maintenance System................................ 49-110 49-00 APU Operation......................................................... 49-112 General Description ..................................................... 49-112 Controls and Indications .............................................. 49-114 Flight Deck Controls................................................ 49-114 External Controls..................................................... 49-116 Synoptic Page ......................................................... 49-118 Operation .................................................................... 49-120 APU Start ................................................................ 49-120 Normal APU Shutdown ........................................... 49-122
TECHNICAL TRAINING MANUAL
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49-ii
CS130_M5_49_APUTOC.fm
49 - Airborne Auxiliary Power
Emergency APU Shutdown .....................................49-124 Monitoring and Test .....................................................49-126 CAS Messages ........................................................49-127
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-iii
CS130_M5_49_APUTOC.fm
49 - Airborne Auxiliary Power
List of Figures Figure 1: APU Left Side View ...............................................49-3 Figure 2: APU Right Side View.............................................49-5 Figure 3: APU Gearbox ........................................................49-7 Figure 4: APU Enclosure and Installation.............................49-9 Figure 5: APU Access Doors..............................................49-11 Figure 6: APU Access Door Struts .....................................49-13 Figure 7: APU Drains..........................................................49-15 Figure 8: APU Mounts and Struts.......................................49-17 Figure 9: APU Eductor........................................................49-19 Figure 10: APU Exhaust .......................................................49-21 Figure 11: APU Drains Schematic........................................49-23 Figure 12: APU Air Intake.....................................................49-25 Figure 13: APU Air Inlet Door Operation ..............................49-27 Figure 14: APU Air Intake Components ...............................49-29 Figure 15: APU Door Indication............................................49-31 Figure 16: APU Inlet Door Operation....................................49-33 Figure 17: APU Inlet Door Mechanical Rigging ....................49-37 Figure 18: Onboard Maintenance System Inlet Door Rigging................................................49-39 Figure 19: APU Door Lockout...............................................49-41 Figure 20: Onboard Maintenance System APU Door Inhibit..................................................49-43 Figure 21: APU Oil System...................................................49-45 Figure 22: APU Oil Components ..........................................49-47 Figure 23: APU Gearbox Oil Components ...........................49-49 Figure 24: APU Lubricating Module......................................49-51 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 25: APU Status Page Oil Indications ........................ 49-53 Figure 26: APU Oil System Schematic ................................ 49-55 Figure 27: APU Oil Servicing ............................................... 49-59 Figure 28: APU Fuel System ............................................... 49-61 Figure 29: APU Fuel Control Unit, Fuel Manifold, Flow Divider, Fuel Nozzles, and Plenum Drain Valve ............................................ 49-63 Figure 30: Fuel Nozzles ....................................................... 49-65 Figure 31: APU Fuel Control Unit Schematic....................... 49-67 Figure 32: APU Fuel Flow Control ....................................... 49-69 Figure 33: Fuel Low Flow Indication .................................... 49-71 Figure 34: APU Starting and Ignition ................................... 49-73 Figure 35: APU Starting and Ignition Starter Motor, Ignition Unit, Igniter Plug, and Igniter Lead......... 49-75 Figure 36: APU Start Power................................................. 49-77 Figure 37: APU Air System .................................................. 49-79 Figure 38: Air System Surge Control Valve, Bleed Air Valve, Inlet Guide Vane Actuator, and Flow Sensors ............................................... 49-81 Figure 39: APU Air Interface ................................................ 49-83 Figure 40: APU Control........................................................ 49-89 Figure 41: APU ECU Control Functions............................... 49-91 Figure 42: APU Start Sequence........................................... 49-93 Figure 43: APU Shutdown Sequence .................................. 49-95 Figure 44: Data Memory Module, Speed Sensor, and EGT Thermocouples.................................... 49-97 Figure 45: APU Electronic Control Unit Location ................. 49-99
TECHNICAL TRAINING MANUAL
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49-iv
CS130_M5_49_APULOF.fm
49 - Airborne Auxiliary Power
Figure 46: Data Memory Module ........................................49-101 Figure 47: ECU Power........................................................49-103 Figure 48: Electronic Control Unit Interface........................49-105 Figure 49: OMS Fuel Shutoff Valve Test............................49-111 Figure 50: APU Operation ..................................................49-113 Figure 51: APU Flight Deck Controls..................................49-115 Figure 52: APU External Controls ......................................49-117 Figure 53: APU Status Performance Indications ................49-119 Figure 54: APU Start ..........................................................49-121 Figure 55: APU Normal Shutdown .....................................49-123 Figure 56: APU Emergency Shutdown...............................49-125
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-v
CS130_M5_49_APULOF.fm
AIRBORNE AUXILIARY POWER - CHAPTER BREAKDOWN AIRBORNE AUXILIARY POWER CHAPTER BREAKDOWN
Starting and Ignition
Construction
1
5
Air Systems
Installation
2
6
Control
Oil
3
7
Operation
Fuel
4
8
49 - Airborne Auxiliary Power 49-00 APU Construction
49-00 APU CONSTRUCTION GENERAL DESCRIPTION The auxiliary power unit (APU) design consists of two modules: the power section and load compressor, and the gearbox with accessories. LEFT SIDE VIEW The following major components are installed on the APU left side: •
Fuel control unit (FCU)
•
Lube module
•
APU drains
•
Ignition unit
•
Thermostatic pressure-relief valve
•
Air/oil cooler (AOC)
The wiring harness provides an interface between the APU and the electronic control unit (ECU). The starter motor and APU generator have their own wiring harnesses to interface with the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-2
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
Load Compressor
Thermostatic Pressure-Relief Valve
Power Section
Air/Oil Cooler
Gearbox
APU Wiring Harness Connectors
Fuel Control Unit Lube Module
CS1_CS3_4900_004
Ignition Unit
Drains
Figure 1: APU Left Side View Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-3
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
RIGHT SIDE VIEW The following major components are installed on the APU right side: •
APU generator
•
Starter motor
•
Bleed air valve (BAV)
•
Surge control valve
•
Fuel manifold and nozzles
Compartment cooling air is drawn through an eductor that surrounds the exhaust to cool the APU oil and provide compartment ventilation. The bleed air duct provides the interface with the aircraft pneumatic system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-4
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
Fuel Nozzles Eductor
Starter Motor
Exhaust APU Generator
Bleed Air Connection
Surge Control Valve
Bleed Air Valve
CS1_CS3_4900_005 4900 005
Fuel Manifolds
Figure 2: APU Right Side View Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-5
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
ACCESSORY GEARBOX SECTION The accessory gearbox is driven through a quill shaft by the high speed torque of the power section. The gearbox contains a series of spur gears to drive the APU accessories. The following accessories are installed on the gearbox: •
Generator
•
Lubricating module
•
Fuel control unit (FCU)
•
Starter motor
•
Air/oil separator
•
Oil fill port
•
High oil temperature sensor
The gearbox also serves as an oil reservoir for the wet-sump lubrication system. An oil drain plug located near the bottom of the gearbox incorporates a magnetic chip collector that can be removed separately.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-6
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-00 APU Construction
Air/Oil Separator Starter Motor Lubricating Module
Generator Fuel Control Unit
High Oil Temperature Sensor Magnetic Chip Collector
CS1_CS3_4900_006
Oil Fill Port
Figure 3: APU Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-7
CS130_49_L2_APUConstruction_General.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
49-10 APU INSTALLATION GENERAL DESCRIPTION The auxiliary power unit (APU) sits in an enclosure that consists of five titanium firewalls and a pair of hinged composite access doors. The APU is suspended in the enclosure at a 13° nosedown attitude. The enclosure isolates the APU from the rest of the structure and acts as a load path to support the APU. Seals are provided around the firewall and door interface. There are three access panels on both the right side firewall and left side firewall and two access panels on the top firewall, which are used to gain access to the system components and structures. There is also one access panel located on the aft firewall, to facilitate the removal of the muffler.
COMPONENT LOCATION The APU installation consists of the following components: •
Access doors
•
APU drains
•
Engine mounts and struts
•
Eductor
•
Exhaust
APU compartment cooling and ventilation is provided by the APU air inlet duct when the APU inlet door opens during APU operation. A pair of hinged access doors provides access for servicing and maintenance of the APU. A drain allows accumulated fluids to exit through a small opening in the left hand access door. The electrical wiring harnesses provide the interfaces between the electronic control unit (ECU) and the APU. The APU wiring harness extends from the firewall connector to the front of the APU, and connect all electrical components. APU harness connectors are threaded, stainless steel, and self-locking. A bleed air duct connects the APU pneumatic output to the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-8
CS130_49_L2_APUInstall_General.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Compartment Cooling Air Inlet
Engine Mount APU Drains
Bleed Air Duct
Drain Line Access Panels (8x)
Exhaust
Eductor Struts
CS1_CS3_4911_003
Hold Open Strut (4x)
Figure 4: APU Enclosure and Installation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-9
CS130_49_L2_APUInstall_General.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
COMPONENT INFORMATION ACCESS DOORS The APU doors are attached to the aircraft structure by goose neck hinges. Fireproof door seals located at the interface of the APU firewall and access doors help to prevent fire propagation and provide containment of Halon when extinguisher is activated. The doors are secured closed by two transverse hook latches and four pin latches. A door seal runs the length of the door interface. There are drain holes along the forward edge of the doors. Accumulated fluid from the APU drains exit through a hole near the forward transverse latch.
CAUTION Do not keep the doors in the full open position if winds are more than 65 kt (120 kph). Damage to door structures, hinges, and supports can occur.
NOTE Open the left door first if the two doors are to be opened. Close the right door first when the two doors are open.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-10
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Pin Latch Closed
Pin Latch Open
Drain Hole
Latch Keeper
Hook Latch
Pin Latch Closed
Drain Hole
Pin Latch Open
CS1_CS3_4913_001
Hook Latch
Figure 5: APU Access Doors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-11
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
ACCESS DOOR STRUTS Four struts hold the access doors open. The struts are fixed to the doors and enclosure. When the doors are opened or closed, the struts fold and stow in clips mounted on the doors. A sleeve locks the strut in the extended position when the doors are opened. Each strut has an adjustable rod end to ensure the door is flush with the fuselage.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-12
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
CS1_CS3_4913_002
Sleeve (4x) Strut
(4x)
Figure 6: APU Access Door Struts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-13
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
APU DRAINS The overboard drains are used to drain the APU inlet plenum and exhaust. The drains aid in detecting fuel leaks at the fuel control unit, inlet guide vane actuator and surge control valve, and oil leaks at the load compressor seal. The last drain allows fluid accumulation in the muffler to drain as well as fuel from an unsuccessful start to drain from the combustor plenum drain orifice. The APU Inlet plenum drain tube disposes of the fluids that enter the APU inlet. The fluids drain on the interior of the APU compartment door near a drain hole. All of the APU drains, except for the inlet plenum drain, are grouped at a drain manifold that is supported off the APU structure. The drain manifold provides sealing via a silicone-rubber bathtub seal to the compartment access door. From this manifold, the system drains into a cavity created by the seal and the APU compartment access door. Fluids drain overboard through a drain hole in the compartment access door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-14
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Surge Control Valve and Inlet Guide Vane Actuator
Load Compressor Seal Witness Drain
Combustor Plenum Drain Orifice/ APU Exhaust
CS1_CS3_4912_001
Fuel Control Unit
Inlet Plenum
Figure 7: APU Drains Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-15
CS130_49_L2_APUInstall_ComponentInformation.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
DETAILED COMPONENT INFORMATION MOUNTS AND STRUTS The APU is suspended in the enclosure at a 13° nosedown attitude by a series of struts and isolators. The system includes four groups of struts that connect to the three mounting points on the APU. The four groups of struts have a suspension system with seven active struts and one catcher strut. A vibration isolator connects each group of struts to the mounting points. The mount system consists of a right side forward isolator with two struts, a left side forward catcher with an adjustable strut, a right side aft isolator with three fixed struts, and a left side aft isolator with one fixed strut, and one adjustable strut. The mount system provides a redundant load path to maintain the APU in position, and provides redundant vibration and shock isolation from the APU to the aircraft. The mount system maintains the APU in position even if one strut fails, or an entire mount is damaged by fire. All isolators react to vertical and lateral loads. The aft isolators also react to thrust loads. The forward left link only reacts to loads in failed condition.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-16
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Strut (7x) Engine Mounts and Isolators (3x)
CS1 CS3 4911 002 CS1_CS3_4911_002
Catcher Strut
Figure 8: APU Mounts and Struts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-17
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
EDUCTOR The eductor provides APU compartment cooling and oil cooling. The eductor uses APU exhaust gas to passively create APU compartment cooling flow. The kinetic energy of the high-velocity exhaust gas draws the flow of lower velocity compartment cooling airstream in a mixing duct. Airflow from the surge system is also introduced in the eductor. Airflow from the compartment is pulled across the oil cooler prior to entering the exhaust stream. The mixed exhaust flow results in lower exhaust system surface temperatures.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-18
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Compartment Cooling Air
Air/Oil Cooler
Eductor
Surge Bleed
CS1_CS3_4921_002
Exhaust
Figure 9: APU Eductor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-19
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
EXHAUST The APU exhaust system is located in the tailcone of the aircraft, aft of the APU compartment. The exhaust has a muffler for noise reduction. The APU exhaust duct attaches to the engine using flexible bellows. The bellows take up the relative motion between the engine and exhaust. A bellows liner provides a smooth surface for the gases to flow through. The bellows bolts to the muffler at the aft firewall. The muffler consists of an outer shell and an inner liner separated by baffles. The inner liner can be removed separately from the outer shell. The muffler is covered by a thermal blanket. The muffler has two lifting points on top for removal. The front muffler face has a fire seal and mounting tabs to mount the exhaust duct to the collar frame at the aft firewall. The muffler has a forward facing drain that expels any fluid accumulation. The drain connects to a flexible tube that drains overboard through the APU access doors The tailpipe is attached to the muffler. The tailpipe is covered by an upper and lower thermal blanket, where it extends through a cutout in the tailcone.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-20
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Upper Thermal Blanket
Tailpipe
Lifting Points
Lower Thermal Blanket
Attachment Point
Liner Muffler
Drain
CS1_CS3_4921_001
Bellows
Figure 10: APU Exhaust Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-21
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
APU DRAINS SCHEMATIC The forward drain disposes of the fluids that enter the APU inlet. The second drain detects degraded seal performance in the fuel control unit (FCU), inlet guide vane (IGVA) actuator, and surge control valve (SCV) actuator. The third drain detects degraded load compressor main shaft seal performance. The aft drain provides a means for disposing of excess fuel in the combustor after an aborted start. It also serves as a drain for any fluid accumulation in the APU exhaust. During normal operations, no fuel is discharged from the aft drain. Telltale drains help find fuel seal failures. They are located in the drain lines from the FCU, IGVA, and SCV. A cap at each drain can be removed and checked for the presence of fuel as part of the troubleshooting procedure for isolating a fuel leak.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-22
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-10 APU Installation
Inlet Guide Vane Actuator
FUEL CONTROL
SURGE CONTROL VALVE Inlet Plenum
Inlet Guide Vane Actuator Telltale Witness Drain
Combustor Case/Educator
Fuel Control Unit
Fuel Control Telltale Witness Drain
Drain Mast Surge Control Valve and Inlet Guide Vane Actuator
Load Compressor Seal Witness Drain
CS1_CS3_4912_002
Surge Control Valve Telltale Witness Drain
Figure 11: APU Drains Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-23
CS130_49_L3_APUInstall_DetailedComponentInfo.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
49-12 APU AIR INTAKE GENERAL DESCRIPTION AIR INLET The auxiliary power unit (APU) air inlet supplies air for APU operation, cooling, and ventilation. The APU inlet door directs air into the air inlet duct. Cooling air is ducted to the APU compartment by splitting the APU inlet air at the inlet door on the aircraft skin. The inlet duct provides an airflow path to the APU plenum. The compartment cooling air ducting provides cooling air to the APU enclosure. The air flows into the APU compartment, circulates around the APU to vent toxic and combustible gases, before exiting through the exhaust-mounted oil cooler and eductor. During ground operation, the velocity of the hot gas exhaust discharge of the APU through the eductor creates a low-pressure around the APU, pulling the cooling air through the system. During flight operation, ram air at the inlet door provides additional force to move the cooling air. APU AIR INLET DOOR The APU inlet door is an outward opening door that scoops air within the boundary layer and directs the air flow to the APU compressor and oil cooler inlet. In the closed position, the door prevents the APU from windmilling in flight.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-24
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU Air Inlet Door
Compartment Cooling Air Duct
CS1_CS3_4912_003
APU Air Inlet Duct
Figure 12: APU Air Intake Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-25
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU AIR INLET DOOR OPERATION The door opens to 45° for ground operations and 30° in flight. When the door is open, an air gap between the inlet door and the fuselage reduce inlet losses, improve flow stability, and reduce door buffeting. The inlet door actuator moves the lever arm to open and close the inlet door. The inlet door has a mechanical stop that limits the door travel to 50°. The APU electronic control unit (ECU) controls the door position. The actuator motor is powered from the 28 VDC EMER BUS. An air inlet door position sensor provides door positioning information to the ECU for door control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-26
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
45° 30°
DC EMER BUS APU Inlet Door Position Sensor
APU OFF
RUN
START
APU ELECTRONIC CONTROL UNIT
APU Door Actuator
LL PU N R TU
LEGEND APU PANEL
ARINC 429 WOW
CS1_CS3_4912_009
Air/Gnd
Figure 13: APU Air Inlet Door Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-27
CS130_49_L2_APUAI_General.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
COMPONENT LOCATION The following components are part of the APU air intake: •
Air inlet door
•
Air inlet door actuator
•
Air inlet door position sensor
AIR INLET DOOR The air inlet door is located on the right side of the vertical stabilizer above the APU compartment. AIR INLET DOOR ACTUATOR The air inlet door actuator is installed on the left side of the air inlet door. AIR INLET DOOR POSITION SENSOR The air inlet door position sensor is installed on the right side of the APU air inlet door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-28
CS130_49_L2_APUAI_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU Air Inlet Door
Air Inlet Door Position Sensor
CS1_CS3_4915_005
Air Inlet Door Actuator
Figure 14: APU Air Intake Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-29
CS130_49_L2_APUAI_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
CONTROLS AND INDICATIONS The APU inlet door status is shown on the STATUS page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-30
CS130_49_L2_APUAI_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
15 °C 15 °C
TAT SAT
ENGINE
103 113 21.8
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L)
APU DOOR Symbol
APU RPM EGT DOOR
100 % 312 °C OPEN
OIL TEMP OIL PRESS OIL QTY
75 °C NORM NORM
OPEN CLOSED OPEN
Condition APU door open. APU door closed. APU door failed open.
TIRE PRESSURE (PSI)
150 224
CLOSED
150
224
224
224
01
01
èè
APU door failed closed. APU door position invalid.
00
00
TEMP
CS1_CS3_4915_006
BRAKE
SYNOPTIC PAGE - STATUS
Figure 15: APU Door Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-31
CS130_49_L2_APUAI_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
DETAILED DESCRIPTION When the ECU receives the APU switch START command, the APU inlet door opens. The ECU supplies power to energize the APU relay and excitation for the air inlet door position sensor. The 28 VDC power is removed from the actuator when the air inlet door position sensor reports the door is in the correct position, based on weight-on-wheels (WOW) sensing. The actuator is driven by the electronic control unit (ECU) logic to the correct position for starting and operating conditions. There are three inlet positions: closed, 30° open for flight, and 45° open for ground operation. If the ECU does not determine that the inlet door is open within 30 seconds after issuing the command, APU starting is inhibited. The ECU supplies 28 VDC to close the inlet door during the normal APU shutdown sequence as well as protective shutdowns. The 28 VDC power is removed when the ECU detects that the inlet door is in the closed position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_49_L3_APUAI_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
EPC 1 APU Relay
INLET DOOR ACTUATOR APU Inlet Door Pwr DC EMER BUS
10
APU ECU Sec DC EMER BUS
10
28 VDC APU ECU
Ground
28 VDC Ground
Open
Close
DC ESS BUS 2
APU OFF
RUN
START
DMC 1 and DMC 2
Start Command Excitation
PULL TURN
INLET DOOR POSITION SENSOR
Feedback
LEGEND
Air/Gnd WOW
EPC
ARINC 429 External Power Center
CS1_CS3_4912_005
APU PANEL
Figure 16: APU Inlet Door Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-33
CS130_49_L3_APUAI_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
MONITORING AND TESTS The following page displays the crew alerting system (CAS) and INFO messages for the APU air intake system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-34
CS130_49_L2_APUAI_MonitoringTests.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
CAS MESSAGES Table 1: CAUTION Message MESSAGE APU DOOR OPEN
LOGIC APU door failed to close when required.
Table 2: INFO Message MESSAGE
LOGIC
49 APU FAULT - APU DOOR APU door failed to open or APU door closed INOP after open or WOW and APU door RVDT fault.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49 - Airborne Auxiliary Power 49-12 APU Air Intake
PRACTICAL ASPECTS CAUTION
APU INLET DOOR MECHANICAL RIGGING The door rigging procedure ensures the door is in the proper closed position in respect to the fuselage and door seal, as well as providing the correct door open positions for ground and flight. Always refer to the Aircraft Maintenance Publication (AMP) for the latest information.
Use caution when manually driving the actuator. A torque of more than 5 in-lb indicates that the actuator has hit its internal stops. Torque in excess of 10 in-lb may damage the APU inlet door actuator, APU inlet door seal or APU inlet door mechanism. Never exceed this torque.
Access to the APU inlet door actuator is through the forward upper access panel on the APU enclosure. The bleed air duct between the bleed air valve (BAV) and the forward firewall must be removed first for access. The door can be manually driven using the 1/4 in. hex manual drive on the door actuator. Turning the manual drive in a clockwise direction to retracts the actuator and opens the door. Turning the manual drive counterclockwise extends the actuator and closes the door. To mechanically rig the door: •
Disconnect the actuator from the inlet door lever arm
•
Manually move the inlet door from fully open to fully closed and ensure the inlet door moves freely
•
Reconnect the lever arm to the inlet door and manually drive the actuator counterclockwise to close the inlet door until it contacts the door seal
•
Manually drive the actuator clockwise to check that the inlet door opens without interference from any components. Manually drive the actuator to close the inlet door.
•
Measure between the inlet door and aircraft skin to make sure a gap of approximately 1 cm (0.4 in) exists.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
Aircraft Skin Inlet Door
2
Door Seal 1
Rod End
1
A torque of more than 5 in-lb indicates that the actuator has hit its internal stops. Torque in excess of 10 in-lb may damage the APU inlet door actuator, APU inlet door seal or APU inlet door mechanism. Never exceed this torque.
2
Clearance between inlet door and fuselage should be 1 cm (0.4 In.)
CS1_CS3_4912_006
NOTE
Manual Drive
Figure 17: APU Inlet Door Mechanical Rigging Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-37
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
OMS APU INLET DOOR RIGGING The door rigging test can be accessed through the onboard maintenance system (OMS) LRU/System operations APU ECU - DOOR RIGGING page. The electrical door rigging test is accomplished after successful completion of the mechanical door rigging. The test lists preconditions that must be followed. The test can only be carried out with the aircraft on the ground and the APU off. When the test is carried out, the door moves. Ensure that personnel and equipment are clear of the door before starting the test. Follow the test instructions on the screen. The door position can be monitored on the OMS and the STATUS synoptic page, and should indicate 45° when open, and 0° when closed. Unsatisfactory results require further mechanical rigging.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
MAINT MENU
RETURN TO LRU/SYS OPS
APU ECU-APU INLET DOOR RIGGING APU INLET DOOR RIGGING
1/10 PRECONDITIONS 1 - SET `APU RUN SW' TO OFF (APU SHUTDOWN) 2 - VERIFY THAT WEIGHT ON WHEELS IS TRUE 3 - DISCONNECT CIRCUIT BREAKER `APU INLET DOOR PWR' 4 - VERIFY `APU DOOR' AND `APU ECU SEC' CIRCUIT 5 - IF THE DOOR IS INHIBITED `OPEN' OR `CLOSED' REFER TO AMP XX-XX-XX TO DEACTIVATE IT RIGGING INHIBIT/INTERLOCK CONDITIONS APU SHUTDOWN: WEIGHT ON WHEELS: INHIBITS OPEN STATUS: INHIBITS CLOSED STATUS:
YYYYY YYYYY L277 B23 L277 B24
PRESS `CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_4912_007
PRESS `CONTINUE' BUTTON TO GO TO NEXT PAGE
Figure 18: Onboard Maintenance System Inlet Door Rigging Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-39
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
APU DOOR LOCKOUT The APU inlet door lockout pin supports the dispatch of the aircraft with an inoperative APU inlet door actuator or position sensor. The door can be manually driven to the open or closed position from the actuator. The lockout pin is inserted into the actuator arm to lock the door at 0º to keep the door closed if the APU is not going to be operated, or at 30º if the APU is required for flight. The APU can be dispatched for flight with the inlet door in either position. If the door is in the open position, the APU must be operating in flight to prevent the APU from windmilling. If the APU is not operated in flight, the airspeed is limited to 250 KIAS. When the door is locked out in the desired position, the onboard maintenance system (OMS) APU door inhibit is applied.
NOTE The APU can operate with full electrical and pneumatic load on the ground with the door open at 30°.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-40
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
$38 /RFNRXW3LQ
Û&ORVHG 'RRU3RVLWLRQ
Û2SHQ 'RRU3RVLWLRQ
CS1_CS3_4915_004
/RFN3LQ+ROH IRU3RVLWLRQ
Figure 19: APU Door Lockout Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-41
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
ONBOARD MAINTENANCE SYSTEM APU DOOR INHIBIT The onboard maintenance system (OMS) LRU/System Operations APU ECU-Inhibit Tests page is used to inhibit the APU inlet door in the open or closed position when the APU inlet door actuator or position sensor is inoperative. When the APU inlet door has been moved to the desired position, select the applicable door inhibit and then select the Continue soft key. Wait for the TEST OK and either the DOOR SUCCESSFULLY INHIBITED OPEN or the DOOR SUCCESSFULLY INHIBITED CLOSED messages to be displayed. Reactivation of the APU Inlet Door To reactivate the APU inlet door, remove the lock pin from the APU inlet door. If the APU inlet door is not closed, manually drive the door to the fully closed position. On the APU ECU-Inhibit Tests page, select APU DOOR INHIBIT RESET. Confirm the lock pin is removed and select the Continue soft key. Stay on the INHIBIT RESET PAGE for a minimum of five minutes. It is possible that a No response from LRU message may appear. After a successful reset, TEST OK and DOOR INHIBITS SUCCESSFULLY REMOVED messages are displayed. Failure to remain on the INHIBIT RESET PAGE for a minimum of five minutes can result in a failed reset sequence. When the reset is complete, do the OMS APU inlet door rigging as per the aircraft maintenance publication (AMP) to confirm the APU inlet door operation.
CAUTION When the APU inlet door is in the APU DOOR INHIBIT OPEN mode in the OMS, make sure that the door is physically open before an APU start. If the APU is started when the APU inlet door is physically closed while in the APU DOOR INHIBIT OPEN mode, damage to the APU can occur. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-42
CS130_49_L2_APUAI_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-12 APU Air Intake
MAINT MENU
RETURN TO LRU/SYS OPS
MAINT MENU
RETURN TO LRU/SYS OPS
APU ECU-DOOR INHIBITS
APU ECU-INHIBIT TESTS
APU DOOR INHIBIT SELECTION
APU DOOR INHIBIT SELECTION
1/5 *** WARNING *** DO NOT LEAVE THE WRENCH ON THE MANUAL DRIVE. UPON STARTUP, IT COULD SPIN OR BE EJECTED AND INJURE
TEST OK DOOR SUCCESSFULLY INHIBITED OPEN
2/5 SELECT FROM THE FOLLOWING ITEMS:
APU Door Inhibit Open *** NOTE *** WHEN USING THE INLET DOOR ACTUATOR MANUAL CRANK, DO NOT APPLY MORE THAN 10 IN-LBS TO THE 1/4 INCH HEX ON THE BACK OF THE ACTUATOR
APU Door Inhibit Closed
PRECONDITIONS TEST OK
1 - SET `APU RUN SW' TO OFF (APU SHUTDOWN)
APU Door Inhibit Reset
2 - VERIFY THAT WEIGHT ON WHEELS IS TRUE
DOOR SUCCESSFULLY INHIBITED CLOSED
3 - DISCONNECT APU DOOR ACTUATOR POWER TEST INHIBIT/INTERLOCK CONDITIONS APU SHUTDOWN: WEIGHT ON WHEELS: INHIBITS OPEN STATUS: INHIBITS CLOSED STATUS:
INHIBITS/INTERLOCK CONDITION
YYYYY YYYYY L277 B23 L277 B24
- INHIBIT STATUS:
PRESS `CONTINUE' BUTTON TO GO TO NEXT PAGE
NONE
PRESS `CONTINUE' BUTTON TO GO TO CLOSE OUT PAGE
TEST OK
PRESS `CANCEL' BUTTON TO GO TO MAIN MENU DOOR SUCCESSFULLY INHIBITED REMOVED
Write Maintenance Reports
Cancel
Continue
Cancel CS1_CS3_4915_003
Continue
Figure 20: Onboard Maintenance System APU Door Inhibit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-43
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49 - Airborne Auxiliary Power 49-90 APU Oil System
49-90 APU OIL SYSTEM GENERAL DESCRIPTION Oil is drawn from the gearbox through a protective screen into the supply pump. After being discharged from the pump, the oil is regulated to maintain a constant pressure throughout auxiliary power unit (APU) operation. The oil then flows to the thermostatic pressure-relief valve (PRV) and on to the oil cooler. If the oil temperature is low or the oil pressure is high, the oil bypasses the cooler and is distributed to the engine and components. The oil is filtered and distributed to the bearings, gears, and generator. Oil is distributed through internal passages and external lines to the various parts of the APU. The oil is monitored for pressure and temperature. If pressure is too low or temperature is too high, the APU shuts down. The generator is scavenged by elements in the lube module. Oil returning from the generator first passes through inline screens. The pumps discharge oil into a generator scavenge filter and empties it into the reservoir. The generator scavenge and lube module filters have bypass switches that cause the APU to shut down if the filter becomes restricted. Oil from the forward bearings and gearbox returns to the reservoir by gravity. The turbine bearing area is scavenged by a single pump element. Oil passes through an inlet screen, through the pump and into the reservoir. Air and oil are separated by an air/oil separator. The reservoir is vented to the exhaust section which prevents overpressurization of the reservoir during APU operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-44
CS130_49_L2_APUOS_General.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
GEARBOX OIL COOLER STARTER
Air/Oil Separator
EXHAUST
GENERATOR Inlet Screen
LUBE MODULE
Scavenge Return to Gearbox
Oil Level
Inlet Screen
THERMOSTATIC PRESSURERELIEF VALVE
Oil Reservoir
Pressure Scavenge Suction Vent Air
CS1_CS3_4991_003
LEGEND Scavenge Pump Return Screen
Figure 21: APU Oil System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-45
CS130_49_L2_APUOS_General.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
COMPONENT LOCATION
OIL LEVEL SENSOR
The oil system components consist of the following:
The oil level sensor is located on the gearbox, aft of the oil servicing door.
•
Lubricating module
•
Oil cooler
•
Thermostatic pressure-relief valve
•
Low oil pressure (LOP) switch
•
Oil level sensor
•
Gearbox
•
Magnetic chip collector
•
High oil temperature sensor
•
Air/oil separator
LUBRICATING MODULE The lube module is installed on the gearbox. OIL COOLER The oil cooler is an air/oil heat exchanger mounted on the left side of the eductor. THERMOSTATIC/PRESSURE-RELIEF VALVE The thermostatic pressure-relief valve is located on the left side of the gearbox. LOW OIL PRESSURE SWITCH The low oil pressure (LOP) switch is located on the front face of the gearbox next to the lubricating module.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-46
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Oil Cooler
Lubricating Module
Thermostatic Pressure-Relief Valve
Low Oil Pressure Switch
CS1_CS3_4991_002
Oil Level Sensor
Figure 22: APU Oil Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-47
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
GEARBOX The gearbox is mounted to the APU load compressor section. The gearbox holds the APU oil supply. MAGNETIC CHIP COLLECTOR The magnetic chip collector is located on the APU gearbox drain plug. OIL TEMPERATURE SENSOR An oil temperature sensor mounted on the gearbox next to the magnetic chip collector. AIR/OIL SEPARATOR The air/oil separator is located right above and to the left of the lubrication module.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-48
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Air/Oil Separator
Drain Plug With Magnetic Chip Collector Gearbox
Oil Temperature Sensor
CS1_CS3_4991_005
Lubricating Module
Figure 23: APU Gearbox Oil Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-49
CS130_49_L2_APUOS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
DETAILED COMPONENT INFORMATION LUBRICATING MODULE The lubricating module is a self-contained unit that provides lubrication and scavenge functions to the generator, gearbox, and main shaft bearings.
Each filter incorporates a filter bypass valve and filter bypass switch to indicate when filter contamination occurs. The filter bypass valve opens when the oil temperature is cold or the filter is clogged. The filter bypass switches signal the ECU when a filter clogs.
The lubricating module incorporates a three-element gerotor pressure pump, a three-element gerotor scavenge pump for clearing oil from the generator, and a single element gerotor scavenge pump for clearing oil from the APU turbine bearing cavity. A pressure regulator maintains a constant lube supply pressure to the engine and generator and a relief valve prevents overpressurization of the oil system. The module incorporates a seal plate to provide sealing features at each oil passageway. This eliminates the need for external tubes at the lubricating module. Deprime Solenoid The deprime solenoid valve is located next to the oil filters on the lubricating module. The deprime solenoid valve opens during cold weather starts to allow air into the oil system. The air helps reduce the drag on the APU during cold start conditions. During APU shutdown, the valve opens to de-oil the lubrication pump. Filters There are two filters installed on the lubricating module. The APU oil filter removes contaminants from the oil before it is distributed to the APU oil system. The a generator scavenge filter removes contaminants from the oil coming from the generator. The oil filter elements are throwaway types contained in a housing that is screwed into the lubricating module housing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-50
CS130_49_L3_APUOS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Lubricating Module
Deprime Solenoid
Seal Plate
Generator Scavenge Filter Element APU Oil Filter Element
CS1_CS3_4991_012
Filter Bypass Switches
Figure 24: APU Lubricating Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-51
CS130_49_L3_APUOS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
CONTROLS AND INDICATIONS The APU oil indications are displayed on the STATUS page. The oil temperature, pressure and quantity are monitored by the ECU. The APU oil pressure shows NORM when the APU is not running.
NOTE The APU oil pressure indication shows NORM when the APU is off. The APU continues to show NORM as the APU starts, and only indicates LOW if there is insufficient oil pressure when the APU is operating above 34% rpm.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUOS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
APU OIL TEMPERATURE Symbol
15 °C 15 °C
TAT SAT
32 °C 144 °C
ENGINE
103 113 21.8
è½½
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L)
Condition APU oil temperature in normal range. Oil temperature at or above high oil temperature amber line threshold. Oil temperature invalid.
APU OIL PRESSURE Symbol
APU EGT DOOR
100 % 312 °C OPEN
OIL TEMP OIL PRESS OIL QTY
75 °C NORM NORM
NORM LOW
Oil pressure below low oil press threshold.
è½
Oil pressure invalid.
TIRE PRESSURE (PSI)
150 224
Oil pressure in normal range.
150
224
224
224
APU OIL QUANTITY Symbol
BRAKE
00
00
TEMP
01
FULL
01
Condition Oil quantity full.
LOW
Oil quantity below low threshold.
NORM
Oil quantity in normal range.
èè½½
Oil quantity invalid.
SYNOPTIC PAGE - STATUS
CS1_CS3_4991_004
RPM
Condition
Figure 25: APU Status Page Oil Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUOS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
DETAILED DESCRIPTION Oil is drawn from the gearbox case reservoir through a coarse inlet screen into the lubrication supply pump elements. The deprime solenoid valve improves APU starting after cold soaked conditions. The deprime solenoid valve is energized to open when the APU start command is received, and the oil temperature is less than -6.6°C (20°F). When the deprime solenoid is opened, air is introduced into the lube pump inlet and the lubrication module loses the ability to circulate oil to the APU. This reduces starter motor drag during cold starting. The valve closes at 60% rpm. On a normal or automatic shutdown, the solenoid is energized at 50% to de-oil the oil system. The oil pressure is regulated to 68 psi. A pressure-relief valve (PRV) opens if the system pressure exceeds 240 psi. The output of the pump is fed to the oil cooler. The thermostatic pressure-relief valve allows the flow of oil to either flow through or bypass the oil cooler, depending on oil temperature. The thermostatic pressure-relief valve allows oil to bypass the oil cooler at temperatures less than 60°C (140°F). Oil flows through the cooler when temperatures are greater than 77°C (170°F). The valve also opens to bypass the oil cooler when there is a differential pressure of 50 psi across the oil cooler assembly. If the thermostatic pressure-relief valve (PRV) fails to close, a high oil temperature shutdown may occur. The oil is filtered and then internally distributed within the gearbox to the oil-cooled generator, engine gears, and bearings. An external line delivers oil to the rear turbine bearing. Both the generator scavenge and APU lube oil filters have filter bypass switches to indicate when filter contamination occurs. On the ground, a contaminated filter causes an APU shutdown and an APU SHUTDOWN advisory message on the engine indication and crew alerting system (EICAS). In flight, an APU caution message appears and the APU does not shut down.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
After lubricating the generator, oil is scavenged from the generator by the generator scavenge elements, and filtered prior to returning to the gearbox. The APU oil drains directly to the gearbox. The turbine bearing compartment oil returns via a dedicated APU scavenge pump element and drains into the gearbox. In scavenging the oil from the bearing cavities, some air is drawn with it since the scavenge pump capacity is greater than the oil flow to the turbine cavity. The scavenged air/oil mixture must be separated and the gear case is vented to prevent the buildup of pressure. The air/oil separator removes the air from the oil as it returns to the gearbox reservoir. The air vents overboard through a line going back to the APU exhaust duct. An oil level sensor, located inside the gearbox adjacent to the oil level sight glass is used to check the APU oil level when the APU is not running. A low oil pressure (LOP) switch monitors the oil flow to the engine bearings. The LOP switch provides protection against low oil pressure conditions. The LOP is a normally closed switch that is checked during prestart and self-test ECU modes. During a prestart built-in test (BIT), the LOP switch has failed if the APU speed is less than 7% and the switch is open. If the LOP switch has failed during the prestart BIT, the APU can start and run but has no LOP protection. An oil pressure auto-shutdown occurs if the APU is on speed, and the oil pressure is less than 35±5 psi for 20 seconds. The oil temperature sensor provides indication, high oil temperature protection, and activation of the deprime solenoid. The normal operating oil temperature is approximately 96°C (205°F) on a standard sea level day. A high oil temperature shutdown occurs when the APU is on speed and the oil temperature is greater than 143°C (290°F) for 10 seconds. A magnetic chip collector collects debris as the oil returns to the gearbox.
TECHNICAL TRAINING MANUAL
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CS130_49_L3_APUOS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Filter Filter Bypass Bypass Switch Switch Filter Bypass Valve
Thermostatic PressureRelief Valve
PressureRegulating Relief Valve
LUBE FILTER
START MOTOR
GENERATOR SCAVENGE FILTER
Turbine Bearing Compartment
Low Oil Pressure Switch
Air Oil Separator Clutch
OIL R COOLE
GENERATOR
Turbine Scavenge Element FCU FWD Bearing Compartment Lubricating Pump
Return Screen
Gen. Scavenge Elements (3x) GEARBOX
Oil Level Sensor Oil Temperature Sensor
Gearbox Air
LEGEND Oil Level Sight Glass
Magnetic Chip Collector Inlet Screen
Regulated Pressure Suction Gearbox Air Scavenge Pressure
CS1_CS3_4991_009
Deprime Solenoid Valve LUBRICATION MODULE
Figure 26: APU Oil System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-55
CS130_49_L3_APUOS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the APU oil system.
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49-56
CS130_49_L2_APUOS_MonitoringTests.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
CAS MESSAGES Table 3: Advisory Message MESSAGE APU OIL LO QTY
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC APU oil low quantity detected.
TECHNICAL TRAINING MANUAL
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CS130_49_L2_APUOS_MonitoringTests.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
PRACTICAL ASPECTS OIL SERVICING The gearbox oil supply provides 700 hours of APU operation without oil servicing. The gearbox sump capacity is 7.32 L (7.74 qt). The APU oil servicing is accomplished when the APU access doors are open. The oil quantity can be checked using the sight glass on the side of the gearbox. The sight glass is located on the aft side of the gearbox near the oil service door. The oil level should be checked 30 minutes after shutdown.
NOTE If the oil level is in the ADD mark, oil must be added to the APU gearbox no later than 30 hours of APU operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-58
CS130_49_L2_APUOS_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-90 APU Oil System
Oil Service Door
FULL L
1 QT
1 QT T
ADD
ADD D
APU OFF
APU U OFF F
Oil Level Sensor
CS1_CS3_4991_010 CS3 4991 010
FULL
SIGHT GAUGE
Figure 27: APU Oil Servicing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-59
CS130_49_L2_APUOS_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
49-30 APU FUEL SYSTEM GENERAL DESCRIPTION The aircraft fuel system supplies the auxiliary power unit (APU) with low-pressure fuel. Fuel enters the fuel control unit (FCU) through the fuel shutoff solenoid valve and is filtered, pressurized, and metered. The FCU supplies metered fuel to the APU for combustion, and high-pressure fuel to operate the surge control valve (SCV) and inlet guide vane (IGV) actuator. The metered fuel flows from the FCU to the fuel flow divider when the fuel shutoff solenoid valve in the FCU is commanded open by the electronic control unit (ECU). The flow divider distributes the fuel between primary flow and secondary fuel manifolds. The primary flow is used for initial starting of the APU. The secondary flow augments the primary flow for acceleration and on-speed operating conditions. The flow divider solenoid is controlled by the ECU. Ten fuel nozzles inject fuel into the combustion chamber for APU operation. When the ECU receives an OFF signal and the APU has completed a cool-down cycle, the fuel shutoff solenoid valve closes and the APU shuts down. On the bottom of the outer combustion case is a plenum drain orifice. It drains the excess fuel from the combustion chamber in case of a no start condition.
NOTE The plenum drain valve is always open, whether the APU is operating or not. An arrow on the valve body points in the direction of fuel flow toward the overboard drain line. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-60
CS130_49_L2_APUFS_General.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER From Aircraft Fuel System
APU FEED SHUTOFF VALVE
FUEL CONTROL UNIT SOL
SURGE CONTROL VALVE
Primary Fuel Flow
Secondary Fuel Flow
INLET GUIDE VANE ACTUATOR
FUEL NOZZLES
ECU
Discrete Fuel Flow Divider SOL
Combustor
Flow Divider Solenoid Plenum Drain Orifice
CS1_CS3_4931_002
LEGEND
Figure 28: APU Fuel System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-61
CS130_49_L2_APUFS_General.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
COMPONENT LOCATION The fuel system consists of the following components: •
Fuel control unit
•
Flow divider assembly
•
Fuel manifolds and fuel nozzles
•
Plenum drain orifice
FUEL CONTROL UNIT The fuel control unit (FCU) is attached to the lubricating module by means of a quick-release v-groove clamp, and is driven by an oil lubricated splined drive shaft. FLOW DIVIDER ASSEMBLY The flow divider is mounted on the lower left side of the turbine plenum. FUEL MANIFOLDS AND FUEL NOZZLES The primary and secondary fuel manifolds surround the combustor plenum. PLENUM DRAIN ORIFICE The plenum drain orifice is located at the lowest point on the plenum.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-62
CS130_49_L2_APUFS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
Fuel Manifolds
Flow Divider
Fuel Control Unit
Fuel Filter
CS1_CS3_4931_003
Fuel Nozzles
Plenum Drain Orifice
Figure 29: APU Fuel Control Unit, Fuel Manifold, Flow Divider, Fuel Nozzles, and Plenum Drain Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-63
CS130_49_L2_APUFS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
DETAILED COMPONENT INFORMATION FUEL NOZZLES There are 10 dual-orifice fuel nozzles equally spaced and mounted on the combustor plenum. The primary and secondary fuel manifolds route fuel to the fuel nozzles. The fuel nozzles inject metered fuel into the combustor. The primary and secondary fuel inlets have screens to prevent contaminants from entering the spray tip. A locating pin positions the air shroud on the nozzle to ensure proper installation in the engine. Fuel for ignition and initial acceleration is supplied by the primary nozzles. The secondary nozzles provide the additional fuel flow as required for operation up to 100% rpm.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-64
CS130_49_L3_APUFS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
Secondary Fuel
Primary Fuel
Screens Locating Pin Compressor Discharge Air In
Cooling Air Coo Shroud Shro
FUEL NOZZLE
CS1_CS3_4931_007
Nozzle Tip
Figure 30: Fuel Nozzles Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-65
CS130_49_L3_APUFS_DetailedComponentInformation.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
DETAILED DESCRIPTION FUEL CONTROL UNIT The fuel control unit is attached to the lubrication module by means of a quick release v-band clamp. The FCU is driven by an oil lubricated spline drive shaft. The APU fuel system is automatically controlled by the electronic control unit (ECU). The aircraft fuel system supplies the APU with low-pressure fuel from the left engine feed line. Fuel enters the FCU through the inlet filter. The filters trap normal fuel pump wear debris. The entire pump discharge flow passes through the filter and feeds a gear pump that supplies the actuator pressure regulator and the metering valve through the high-pressure fuel filter.
A temperature sensor signals the ECU to adjust flow rates for starting the APU in cold conditions and at high altitude. A flow meter resolver provides fuel flow feedback to the ECU. Once rated speed is reached, fuel flow is modulated to meet the demands of varying pneumatic and electrical loads. When the ECU receives an OFF signal and the APU has completed a cooldown cycle, power is removed from the fuel shutoff solenoid and the APU shuts down. Whenever a protective shutdown is initiated, the signal to the FCU torque motor is automatically removed as well.
A high pressure-relief valve protects the FCU. The relief valve opens at 1150 psi and also relieves pump pressure on all FCU shutdowns. The pressure regulating valve provides 250 ± 25 psi hydraulic pressure used to activate the inlet guide vane actuator and the surge control valve. The fuel metering valve meters fuel flow using a torque motor. The metered fuel is a function of ECU current signal. An increase in current increases fuel flow. Whenever a protective shutdown is initiated, the current to the torque-motor is automatically removed. At 7% rpm, the fuel shutoff solenoid allows fuel to flow to the flow divider and flow divider solenoid. When the solenoid is de-energized, APU operation is terminated.
NOTE If the fuel shutoff solenoid fails to open at 7% speed, a protective shutdown occurs. A leaking fuel shutoff solenoid that allows fuel into the APU before 7% speed causes torching.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-66
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER
SOL
Fuel Temperature Sensor
Fuel Metering Valve
Actuator Pressure Regulator
High-Pressure Fuel Filter Inlet Guide Vane Actuator
FUEL NOZZLES
HighPressure Relief Valve
Flowmeter Resolver
Combustor
Spline Drive Shaft SOL
Fuel Shutoff Solenoid
Drain Orifice
LUBE MODULE GEARBOX
HighPressure Gear Pump
FUEL CONTROL
Fuel Supply Differential Pressure Regulator
FUEL INLET FILTER ELEMENT
Actuator Return Fuel (Low Pressure) Metered Fuel Pump Discharge High-Pressure Regulated Pressure Muscle Press Flow Divider SOL
Flow Divider Solenoid
CS1_CS3_4931_005
LEGEND SURGE CONTROL VALVE
Figure 31: APU Fuel Control Unit Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-67
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER AND FUEL FLOW DIVIDER SOLENOID The flow divider and flow divider solenoid distribute the fuel between primary flow and secondary fuel manifolds. The flow divider solenoid also improves start capability in cold weather. Fuel flows to the primary manifold for APU light-off and initial acceleration. The ECU uses P2, T2, and speed signals to control the fuel flow divider solenoid valve. During an APU start on the ground or at low altitude, the normally open flow divider solenoid is energized closed until 30% rpm. The solenoid also energizes closed when the APU is operating above 23,000 ft or T2 less than 13°C (55°F). This allows the primary fuel manifold pressure to increase above 125 psi and provide better atomization for increased combustion performance. Fuel flow to the secondary manifold is controlled by the flow divider check valve that opens at 120 psi (25-40% rpm). Incorrect flow divider sequencing results in an APU protective shutdown. The flow divider assembly is a line replaceable unit.
NOTE If flow divider sequencing is incorrect, no acceleration, no flame, overtemperature and underspeed, auto-shutdowns could occur.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-68
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
FUEL FLOW DIVIDER APU RPM < 30% APU Operating Above 23000 ft T2 < 13°c (55°f)
APU ELECTRONIC CONTROL UNIT
From Fuel Control Unit SOL
Solenoid Energized Closed
Flow Divider Check Valve (open at 120 psi (25-40% rpm))
Secondary Fuel Manifold
Primary Fuel Manifold
FUEL NOZZLES
LEGEND
SOL
CS1_CS3_4931_009
Discrete Fuel Flow Divider
Flow Divider Solenoid
Figure 32: APU Fuel Flow Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-69
CS130_49_L3_APUFS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
CONTROLS AND INDICATIONS The APU actual fuel flow is compared to the ECU commanded fuel flow. If a disagreement occurs, the LO FLOW indication on the FUEL synoptic page appears and the filter symbol turns from green to yellow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-70
CS130_49_L2_APUFS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-30 APU Fuel System
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
10000 KG 2500 KG
20 °C 3000 KG
20 °C 3000 KG 4000 KG LO FLOW
SYNOPTIC PAGE - FUEL
CS1_CS3_4931_006
APU
Figure 33: Fuel Low Flow Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-71
CS130_49_L2_APUFS_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
49-40 APU START AND IGNITION SYSTEMS GENERAL DESCRIPTION The auxiliary power unit (APU) starting and ignition system accelerates the APU to the starting speed and ignites the fuel/air mixture in the combustor. The ECU provides a start command to BPCU 2 when the APU panel START switch is moved to START and the APU inlet door has reached the open position. An APU start uses electrical power supplied by TRU 3 via the TRU start contactor (TSC) or BATT 2 via the battery start contactor (BSC). Once a power source is selected, BPCU 2 closes the APU start contactor (ASC) and the APU starter motor drives the APU through a sprag clutch mounted on the gearbox. The APU start request terminates at 50% rpm and power is removed from the starter. The APU continues to accelerate to a self-sustaining speed. The ignition system provides electrical energy to create a spark at the igniter plug for starting the APU. The ECU controls and powers the ignition system. The ignition system is energized once the start request is made and terminates at 60% rpm. The ignition unit delivers 18,000 V to the igniter plug.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-72
CS130_49_L2_APUSIS_General.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
APU Starter
Igniter
BPCU 2
ASC
Ignition Unit
DMC 1 AND DMC 2
APU ECU BSC
TRU 3
BATT 2
DC BUS 2
DC ESS BUS 2
LEGEND ASC BPCU BSC ECU TSC
APU Start Contactor BUS Power Control Unit Battery Start Contactor Electronic Control Unit TRU Start Contactor ARINC 429
CS1_CS3_4941_002
TSC
Figure 34: APU Starting and Ignition Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-73
CS130_49_L2_APUSIS_General.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
COMPONENT LOCATION The starting and ignition system consists of the following components: •
Starter motor
•
Ignition unit
•
Igniter plug and igniter lead
STARTER MOTOR AND CLUTCH ASSEMBLY The starter motor and clutch assembly is installed on the gearbox, above the generator. The DC starter motor has a brush wear indicator pin. The brush wear indicator pin must be seen through the side window of the indicator housing. If any part of the white indicator pin is visible, the brushes are acceptable for continued use. IGNITION UNIT The igniter unit is located on the bottom of the APU, below the oil cooler. IGNITER PLUG AND IGNITER LEAD The igniter plug is installed on the left side of the combustion section. A single air-gap-type igniter provides ignition for the APU. The igniter consists of an center electrode, an insulator, and an outer shell. The outer shell has cooling holes to aid in reducing tip temperatures. An igniter lead connects the igniter to the ignition unit.
CAUTION Voltage produced by the ignition system is lethal. Caution must be observed when working with the system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-74
CS130_49_L2_APUSIS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
Handcrank
Brush Wear Indicator
Igniter Plug
STARTER MOTOR
CS1_CS3_4941_003
Igniter Lead
Ignition Unit
Figure 35: APU Starting and Ignition Starter Motor, Ignition Unit, Igniter Plug, and Igniter Lead Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-75
CS130_49_L2_APUSIS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
The BPCU commands the ASC open and commands CDC 4 to open the TSC when the APU start from TRU 3 is terminated or aborted.
DETAILED DESCRIPTION APU START POWER The APU can be started using transformer rectifier unit (TRU) 3 or BATT 2. A TRU APU start has priority over a battery APU start if there is no TRU failure in the system. The APU starter motor power is determined by the electrical system configuration. The electrical system applies power to the APU starter motor when BPCU 2 receives an APU start request from the APU ECU. BPCU 2 configures the electrical system for start. The ECU monitors the voltage available to the starter. If an open circuit is detected, a maintenance message is generated. The APU start request discrete signal sent by the APU ECU becomes inactive during a normal start when APU speed exceeds 50% or if the ECU detects an aborted start condition. BPCU 2 aborts and inhibits the APU start function if the start request from the ECU is not terminated after 60 seconds. Two failed APU start attempts followed by one successful attempt are permitted during normal operation. Do not perform more than three starts/start attempts in one hour. A two-minute delay must be observed between start attempts. A cooldown period is required for the TRU after three failed starts. The battery must cooldown and recharge after three failed start attempts. If any power contactor failure is detected, or there is an AC power transfer, the APU start is aborted.
BATTERY START An APU start from battery occurs when both batteries are available and the TRU 3 start is inhibited. Two batteries are in use during the APU starting interval. Battery 2 is dedicated to the APU starter, and battery 1 powers the aircraft. An APU start from battery 2 is inhibited when it is the only source of power on the aircraft or any associated APU start power contactors are failed. For a battery start, BPCU 2 commands battery line contactor (BLC) 2 to open. Following confirmation the BLC 2 status is open, BPCU 2 requests CDC 2 to close the BSC. CDC 2 monitors the status of BSC contactor and reports it to BPCU 2. When the BSC is confirmed closed, BPCU2 commands the ASC to close. When the APU start from battery 2 is terminated or aborted, the BPCU commands the ASC open and commands CDC 2 to open the BSC.
TRU START APU starts from TRU 3 is inhibited if any TRU is not powered. For a TRU APU start, BPCU 2 commands CDC 4 via CDC 2 to close the TRU start contactor (TSC). CDC 4 commands the TSC coil to close. CDC 4 monitors and sends the status of the TSC contactor to BPCU 2. Following confirmation the TSC is closed, BPCU 2 closes the APU start contactor (ASC).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-76
CS130_49_L3_APUSIS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-40 APU Start and Ignition Systems
Start Power Selection Logic
EPC 3
TRU 1 TRU 2 TRU 3
TRU
BATT 1 BATT 2
BATT
APU START
TSC TRU 3
AC ESS BUS
TSC Status CDC 4 CDC 4-4-10 TSC Ctrl DC BUS 2
LEGEND ASC BLC BPCU BSC CDC EPC RLC TSC
SSPC 3A RLC
Logic: Advanced
APU Start Contactor Battery Load Contactor BUS Power Control Unit Battery Start Contactor Control and Distribution Cabinet Electrical Power Center Ram Air Turbine Load Contactor TRU Start Contactor ARINC 429 TTP
EPC 2 BSC BATT 2
BATT DIR BUS 2
ASC
APU STARTER MOTOR
BSC Status CDC 2 CDC 2-7-12 BSC Ctrl
APU RUN
START
SSPC 3A
Logic: Advanced PULL TURN
APU ECU
BLC
BPCU 2 ASC Status APU Start Command
APU PANEL
APU Starter Voltage Sensing
CS1_CS3_4941_005
OFF
DC ESS BUS 2
Figure 36: APU Start Power Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-77
CS130_49_L3_APUSIS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
49-50 APU AIR SYSTEMS GENERAL DESCRIPTION The auxiliary power unit (APU) supplies compressed air for air conditioning and main engine starting from the load compressor. The pneumatic output is through the bleed air valve (BAV).
The BAV is a pneumatically actuated butterfly valve used to control bleed flow to the airplane. The BAV is normally in the closed position and uses load compressor bleed air controlled by a solenoid to provide the muscle for valve opening.
The load compressor takes air from the same inlet as the power section compressor. It compresses the air and supplies it to the aircraft through a bleed air valve. The bleed air valve shuts off the air flow from the load compressor when there is no demand placed on the APU by the aircraft pneumatic system. The load compressor provides the maximum amount of air flow that the aircraft demands. Inlet guide vanes (IGVs), located between the load compressor and the air inlet, match the compressor flow to the demand. The inlet guide vanes are positioned by an inlet guide vane actuator (IGVA). The ECU controls the IGVA and surge control valve (SCV), using environmental data from P2 (pressure) and T2 (inlet temperature), load compressor output (Pt and P), and load demand from the aircraft. When there is no demand, the ECU closes the IGVs to the lowest possible setting, so the load compressor imposes a minimum load on the power section. In this condition, the bleed air valve is closed and all air is directed through the SCV. The SCV prevents load compressor surge for all operating modes, altitudes, and inlet temperatures. Surge is prevented by opening the SCV when the aircraft pneumatic system takes little or no airflow, or when the BAV is closed. During the APU shutdown sequence, the ECU initiates pneumatic unloading by closing the IGVs and BAV, and opening the SCV.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-78
CS130_49_L2_APUAS_General.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
T2
Fuel In Out
To ECU P2
IGVA
To ECU
Oil Cooler
Inlet Guide Vanes
ECU
Sensor PT Sensor SOL
To Aircraft Pneumatic System
Surge Bleed Air Exhaust LEGEND
Fuel In Out
SCV
BAV
To ECU
BAV ECU IGVA LVDT PT P2 SCV TM T2 ¨3
Gearbox Vent APU Inlet Air APU Combustion Air Bleed Air Fuel Bleed Air Valve Electronic Control Unit Inlet Guide Vane Actuator Linear Variable Differential Transformer Total Pressure Inlet Pressure Sensor Surge Control Valve Torque Motor Inlet Temperature Sensor Differential Pressure Discrete
CS1_CS3_4951_008
¨P
Air Inlet
Load Compressor Outlet
Figure 37: APU Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-79
CS130_49_L2_APUAS_General.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
COMPONENT LOCATION
INLET GUIDE VANE ACTUATOR
The air system consists of the following components:
The inlet guide vane actuator is located below the load compressor section.
•
Surge control valve
•
P2 sensor
BLEED AIR VALVE
•
Total pressure sensor
•
Differential pressure sensor
The bleed air valve is on the right side of the APU compartment in the bleed air duct.
•
Inlet temperature sensor
•
Inlet guide vane actuator
•
Bleed air valve
SURGE CONTROL VALVE The surge control valve is located in the surge bleed duct on the right side of the APU. P2 SENSOR The P2 sensor is installed on the right side of the inlet plenum. TOTAL PRESSURE SENSOR The total pressure (Pt) sensor is located on the bottom of the APU compressor plenum. DIFFERENTIAL PRESSURE SENSOR The differential pressure (P) sensor is located on the bottom of the APU compressor plenum. INLET TEMPERATURE SENSOR The inlet temperature (T2) sensor is located on the bottom of the APU compressor plenum.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-80
CS130_49_L2_APUAS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
ǻP Sensor Bleed Air Valve
PT Sensor
P2 Sensor Surge Control Valve
LEGEND PT Sensor P2 Sensor T2 Sensor ǻP Sensor
Total Pressure Sensor Inlet Pressure Sensor Inlet Temperature Sensor Differential Pressure Sensor
Inlet Guide Vane Actuator
CS1 CS3 4951 003 CS1_CS3_4951_003
T2 Sensor
Figure 38: Air System Surge Control Valve, Bleed Air Valve, Inlet Guide Vane Actuator, and Flow Sensors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-81
CS130_49_L2_APUAS_ComponentLocation.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
DETAILED DESCRIPTION APU AIR INTERFACE The APU electronic control unit (ECU) communicates with the aircraft systems via the data concentrator module cabinet (DMC) 1 and DMC 2 over ARINC 429 BUSES.
The pneumatic output of the APU is supplied to the aircraft through a bleed air valve. The bleed air valve is a solenoid controlled valve, receiving 28 VDC form the ECU. A microswitch in the valve provides an indication when the valve is open.
The APU generator control unit (GCU) supplies electrical load information to the ECU and receives a ready-to-load signal when the APU is on-speed. The ECU receives environmental control system (ECS) demands from the integrated air system controller (IASC) 1 and IASC 2. The ECS demands are based on air data system information and whether the aircraft is on the ground or in flight. The left electronic engine control (EEC) and right EEC provide main engine start (MES) and MES termination commands. The control of the pneumatic output is based on ECS or MES demands as well as the APU inlet temperature (T2) and inlet pressure (P2). The pneumatic output is based on the position of the inlet guide vanes. The IGVA is positioned by a torque motor driven by the ECU. The IGVA provides position feedback to the ECU. When the pneumatic demand changes, due to ECS pack valve closures or the termination of MES, the surge control valve opens to prevent an APU surge. The surge control valve is positioned by a torque motor driven by the ECU. The ECU receives position feedback from the SCV. The SCV position is based on the load compressor total pressure (Pt) and differential pressure (∆P). The Pt, ∆P, and P2 sensors receive 10 VDC excitation from the ECU and supply outputs proportional to the sensed pressure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-82
CS130_49_L3_APUAS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
AIR COCKPIT
FWD
CABIN
C
AFT
H
FWD
MAN TEMP
CARGO
VENT
OFF
LO HEAT
OFF
HI HEAT
ON
TRIM AIR
PACK FLOW
RECIRC AIR
OFF
HI
OFF
L PACK FAIL OFF
FAIL
IGVA
Command Position
SCV
28 VDC Valve Position
BAV
R PACK XBLEED MAN CLSD
AUTO
MAN OPEN
L BLEED
OFF
Command Position
FAIL OFF R BLEED FAIL
APU BLEED
OFF
FAIL OFF
IASC 1 and IASC 2 APU GCU
LEGEND
DMC 1 and DMC 2
∆P
LEFT EEC and RIGHT EEC
10 VDC Excitation
PT P2
Air Data Information
T2 Air/Gnd
APU ELECTRONIC CONTROL UNIT
WOW
BAV DMC EEC GCU IASC IGVA PT P2 SCV T2 ∆P
Bleed Air Valve Data Concentrator Unit Module Cabinet Electronic Engine Control Generator Control Unit Integrated Air System Controller Inlet Guide Vane Actuator Total Pressure Inlet Pressure Sensor Surge Control Valve Inlet Temperature Sensor Differential Pressure ARINC 429
CS1_CS3_4950_001
AIR PANEL
Figure 39: APU Air Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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49-83
CS130_49_L3_APUAS_DetailedDescription.fm
49 - Airborne Auxiliary Power 49-50 APU Air Systems
APU AIR CONTROL The load compressor output is regulated by variable IGVs located just upstream of the load compressor impeller. Output is based on aircraft environmental control system (ECS), main engine start (MES), and electrical load demands as indicated by aircraft signals to the APU electrical control unit (ECU). The IGV position is a function of P2, T2, and load demand. During an APU start, the IGVs are closed to 15° up to 60% rpm and then opened to 22° if altitude is below 25,000 ft. This removes any pneumatic load from the load compressor. Above 25,000 feet, the IGVs are held at 15°.
The surge control valve (SCV) ducts air from the load compressor to the APU exhaust. The SCV makes sure there is a minimum flow of air through the load compressor. The surge margin set point is the minimum quantity of air that should flow through the load compressor to prevent load compressor surge. The ECU calculates the surge margin set point using: •
Inlet pressure (P2)
•
Inlet temperature (T2)
To minimize fuel consumption, when there is no demand, the ECU closes the IGVs to the lowest possible setting that provides a constant airflow through the compressor during all loading conditions so the load compressor imposes minimum load on the power section. The ECU logic for determining IGV position includes a compensation factor for load compressor deterioration, based on the APU operating hours recorded by the data memory module (DMM).
•
IGV position
•
Bleed mode
•
Air/ground mode
To optimize fuel consumption in the ECS mode, the pneumatic load provided by the APU is matched to the aircraft requirements. The pressure and flow requirements vary with ambient conditions and whether the aircraft is on the ground or in flight.
During MES, the APU ECU opens the IGVs to 90° and modulates the SCV to provide maximum bleed for engine starting. The priority is given to the pneumatic load during engine start. If required, some APU generator load sheds.
As the combined loads of the generator and the load compressor are imposed on the power section, the speed governing function of the ECU responds by increasing fuel flow. Under some extremely high combination loads or with significant deterioration of the power section, the turbine temperature limit is reached.
To prevent damage to air conditioning components as well as overpressure and overtemperature conditions, the APU software logic limits the APU bleed pressure and temperature to: •
65 psi maximum bleed pressure
When a load change results in a temperature higher than its limit, the ECU moves the IGVs toward a more closed position, reducing load compressor airflow, and thus reducing the pneumatic load and decreasing the turbine temperature. The electrical load remains unchanged and has priority over the pneumatic load.
•
246ºC (475ºF) maximum bleed temperature
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The ECU uses total pressure (Pt) and differential pressure (P) to maintain the load compressor output at or above the stall margin set point.
In the event of a failure, the ECU opens the SCV and closes the IGVs. This gives priority to electric operation over bleed air demand in the event of a surge system malfunction.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-50 APU Air Systems
Table 4: APU AIR System Operating Modes MODE Startup
READY-TO-LOAD (RTL) (Generator only)
Environmental Control System
Main Engine Start
ELECTRICAL LOAD AVAILABILITY
DEFINING CONDITIONS •
APU ON signal present
•
MOMENTARY initiates start
START
•
Speed >95% +2 seconds
•
BLEED-AIR COMMAND OFF
•
Not available
PNEUMATIC SYSTEM STATUS •
IGVs positioned engine start
during •
Timed acceleration starting logic applies
•
SCV fully open
•
Mode ends when APU reaches 95% speed
•
Bleed air not available
•
•
IGVs positioned at 15° for altitude >25,000 ft
•
SCV fully open
Bleed air not available, only because APU is not selected as bleed air source at conditions where bleed air would normally be available
•
Bleed air available
•
Electrical load has priority
•
IGVs set to ECS position schedule • by logic in the ECU, as a function of inlet temperature (T2), inlet pressure (P2), A/C configuration, and flow mode
•
Surge control system active
signal
Speed >95% +2 seconds
•
BLEED-AIR COMMAND ON
•
Main engine start is not requested
•
One or both packs on
•
Altitude <23,000 ft
Available
Available
at
15°
•
Speed >95% +2 seconds
•
Bleed air available
•
BLEED-AIR COMMAND ON
•
IGVs set at 90° (full open)
•
Either left or right main engine start signal is true
•
Surge control system active
•
Altitude <23,000 ft
Available
COMMENT
•
If power capability is exceeded, pneumatic load is cut back, as needed, by closing the IGVs
Pneumatic power has priority. If power capability is exceeded, load shed is sent to airplane to reduce electrical load, so that IGVs do not have to be cut back
CS1_CS3_TB49_001
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-85
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49 - Airborne Auxiliary Power 49-50 APU Air Systems
MONITORING AND TEST The following page provides the crew alerting system (CAS) messages for the APU air system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-86
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49 - Airborne Auxiliary Power 49-50 APU Air Systems
CAS MESSAGES Table 5: CAUTION Message MESSAGE APU BLEED FAIL
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC Failure of APU providing bleed air when requested or bleed leak detection fail.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
49-60 APU CONTROL GENERAL DESCRIPTION The auxiliary power unit (APU) electronic control unit (ECU) controls and monitors the APU operation and provides an interface with the aircraft. The APU ECU uses inputs from the exhaust gas temperature (EGT) thermocouples and the speed sensor to control the start and acceleration sequence, speed governing, load sequencing, and the shutdown sequence. The ECU has field-loadable software.
The ECU shuts down the APU when the built-in test (BIT) identifies conditions that can damage the engine or the aircraft for continued APU operation. The ECU also has the authority to inhibit starting of the APU when the BIT indicates APU operation may cause damage.
The EGT thermocouple senses the temperature of the exhaust gas that exits out of the turbine section. The speed sensor provides two independent speed signals to the ECU. The APU ECU operates on 28 VDC power provided by DC ESS BUS 2, with the DC EMER BUS supplying a secondary source of power. The APU is started using the APU switch on the APU panel. The ECU controls the ignition and fuel automatically, based on environmental conditions. The APU runs at a variable speed (91-98%) on the ground, and is governed to 100% rpm in flight. The ECU sends a ready-to-load signals to allow electrical or pneumatic loading when the APU is approaching its governed speed. The data memory module (DMM) is an electronic data storage device mounted on the APU and is used for recording parameters, such as APU operating hours, starts, number of unsuccessful starts, and the APU serial number. The ECU monitors the APU components for faults, recording, and reporting. A failed ECU causes an immediate APU shutdown and an APU SHUTDOWN EICAS advisory message to be displayed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-88
CS130_49_L2_APUControl_General.fm
49 - Airborne Auxiliary Power 49-60 APU Control
DC EMER BUS
APU OFF
RUN
EICAS
DC ESS BUS 2 START
SYN ELECTRONIC
PULL TURN
OMS
CONTROL UNIT
AHMS Speed Sensor
APU PANEL
INLET GUIDE VANE ACTUATOR
Oil Cooler
Exhaust Gas Temperature Thermocouples
DMM GENERATOR
G e a r b o x
IGN UNIT FUEL CONTROL UNIT
DMM IGN
BLEED AIR VALVE
SURGE CONTROL VALVE
Data Memory Module Ignition Unit Discrete Fuel RS-485
APU Fuel Feed Line
CS1_CS3_4961_008
LEGEND
To ACFT Pneumatic System
Figure 40: APU Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-89
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49 - Airborne Auxiliary Power 49-60 APU Control
APU ECU CONTROL FUNCTIONS Prestart Sequence
Normal Shutdown
The ECU performs integrity checks to verify that the APU is safe to start and operate. The requirements for an APU start are:
The ECU initiates APU shutdown by injecting a false overspeed signal into the programmable logic device (PLD), which provides overspeed protection. The ECU uses two independently operating PLD devices. The function of each overspeed PLD is to determine if an APU overspeed condition exists and cause a shutdown.
•
ECU BIT self-test passed
•
ECU BIT test of sensors and controls
•
APU door full open signal
Start Sequence When the inlet door reaches the proper position, the ECU sends a start enable command to BPCU 2 to power the APU starter. The ECU powers the ignition unit, and then energizes the fuel solenoid. As the APU accelerates to operating speed, the ECU controls fuel flow via the fuel control unit torque motor as a function of acceleration, fuel schedule, and EGT. The ignition unit and starter command are deactivated when the speed thresholds are attained. On-Speed Governing Exhaust gas temperature and engine speed are continuously monitored by the ECU. Once on-speed, the ECU maintains engine speed at 100% rpm. On the ground, the APU is governed between 91% and 98% rpm based on environmental conditions. Ready-to-load signals are provided for electrical and pneumatic systems.
The primary mode for shutdown is by removing power from the fuel solenoid. Each PLD is capable of shutting down the APU in the event of an overspeed. In order to limit potential fuel solenoid dormancy, at every normal APU shutdown, the APU fuel shut-off solenoid is exercised, alternating the commanding PLD at each APU shutdown. Protective Shutdowns The ECU shuts down the APU when the BIT identifies conditions that can damage the engine or the aircraft. APU protection is provided by logic schemes that determine whether the APU is safe to operate and protective shutdown logic that shuts down an operating APU. Fault Detection The ECU monitors the APU components for faults, recording and reporting. A failed ECU causes an immediate APU shutdown and an APU SHUTDOWN EICAS advisory message to be displayed. The ECU reports faults to the onboard maintenance system (OMS).
Pneumatic Loading Pneumatic loading for the aircraft environmental control system and main engine start operation are provided by the APU through the ECU control of the IGVs. The ECU positions the IGVs in response to bleed demands. The ECU regulates the APU pneumatic output by sensing the APU EGT and comparing it to a predetermined schedule within the ECU. APU surge is prohibited through the ECU control of the SCV actuator. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-90
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49 - Airborne Auxiliary Power 49-60 APU Control
Prestart Sequence Start Sequence On-speed Governing APU ECU
Pneumatic Loading Normal Shutdown Protective Shutdown Fault Detection
CS1_CS3_4961_009
APU ECU
Figure 41: APU ECU Control Functions Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-91
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49 - Airborne Auxiliary Power 49-60 APU Control
APU START The APU switch signals the ECU when it is selected to the START position. The ECU opens the APU feed shutoff valve and the APU inlet door. When the air inlet door is fully open, the inlet door position sensor sends a door open signal to the ECU. The APU start sequence is as follows: •
0% rpm, the ECU energizes the starter and ignition circuit
•
0% rpm, the IGVs positioned to 15°
•
7% rpm, the fuel solenoid valve opens
•
50% rpm, the starter de-energizes
•
60% rpm, the ignition unit de-energizes
•
On ground, ready to load signal is 3% below governed speed signal (88% rpm to 95% rpm)
•
In flight, ready to load signal is 95% rpm
•
91-98% rpm, governed speed range on ground
•
100% rpm, governed speed in flight
•
106% rpm, overspeed shutdown is initiated
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
APU OFF
RUN
START
106%
Maximum Speed
100% 98% 95%
100% Governed Speed (Flight and Groomer Mode) Ready to Load (Flight)
91%
Electrical and Pneumatic Loading on Ground (Ready to Load 88% to 95%)
PULL TURN
60%
Ignition Unit De-energized
50%
Starter De-energized
7% 0%
Fuel Solenoid Valve Opens Ignition Unit Energized
CS1_CS3_4961_015
• APU switch to START position and release to RUN • APU feed SOV open • Air inlet door open • Starter energized
Figure 42: APU Start Sequence Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-93
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49 - Airborne Auxiliary Power 49-60 APU Control
APU SHUTDOWN
Cooldown Cycle
The ECU controls the APU shutdown. There are two types of shutdown conditions: normal and protective.
The cooldown period prevents oil coke at the turbine bearing and fuel nozzles. During cooldown, the ECU:
Normal Shutdown A normal shutdown is initiated from the APU panel when the APU switch is moved to the OFF position. The normal shutdown sequence is as follows:
•
Removes the bleed air ready-to-load signal
•
Closes the bleed air valve
•
Closes the inlet guide vanes to 22°
•
Opens the surge control valve
•
28 VDC ON signal removed from ECU
•
De-energizes the generator
•
ECU receives the OFF signal
•
Starts the 60 second timer
•
Generator ready to load signal is removed
•
60 second cooldown period starts
•
After cooldown, the ECU tests the overspeed circuit closes the FCU shutoff valve and removes power from the metering valve -
Protective Shutdown A protective shutdown is similar to the normal shutdown but has no cooldown cycle.
ECU tests the overspeed circuit which closes the fuel solenoid shutoff valve and shuts down the APU
NOTE
-
During this time, the ECU tries to maintain the APU speed by modulating the FCU torquemeter current
The APU fuel shutoff valve and air inlet door close for both normal and protective shutdowns.
-
If the APU continues to run, the fuel shutoff solenoid valve has failed to close
-
If the fuel solenoid shutoff valve has failed, the ECU reduces the FCU torquemotor current to 0 to shut down the APU
•
20% speed, the APU inlet door is commanded closed
•
7% rpm, the APU fuel shutoff valve (SOV) closes
•
At less than 7% speed, an APU restart can be initiated
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-94
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49 - Airborne Auxiliary Power 49-60 APU Control
APU Operating
60 Second Cooldown On Speed
100%
APU RUN
START
APU Switch Off
LL PU RN U T
• Ready-to-load signal removed • Bleed air valve closed • Inlet guide vanes closed • Surge control valve open • Generator off • 60 second timer starts for cooldown
20% APU Inlet Door Commanded Closed
20%
7% APU Feed Shutoff Valve Closes
7%
APU Restart Possible 0%
CS1_CS3_4961_018
OFF
Figure 43: APU Shutdown Sequence Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-95
CS130_49_L2_APUControl_General.fm
49 - Airborne Auxiliary Power 49-60 APU Control
COMPONENT LOCATION The control system consists of the following components: •
Exhaust gas temperature thermocouple
•
Speed sensor
•
Data memory module
•
Electronic control unit
EXHAUST GAS TEMPERATURE THERMOCOUPLE The APU consists of two exhaust gas temperature thermocouples probes that are located below the APU turbine casing and extend into the exhaust gas stream inside the APU. SPEED SENSOR The speed sensor is located on the bottom of the load compressor case. DATA MEMORY MODULE The data memory module is installed on the left side of the APU air inlet plenum.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
DMM
Speed Sensor
CS1_CS3_4961_006
EGT Thermocouples
Figure 44: Data Memory Module, Speed Sensor, and EGT Thermocouples Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-97
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49 - Airborne Auxiliary Power 49-60 APU Control
ELECTRONIC CONTROL UNIT The electronic control unit (ECU) is located in the mid equipment bay.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
CS1_CS3_4961_005
MID EQUIPMENT BAY AFT RACK
APU ECU
Figure 45: APU Electronic Control Unit Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-99
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49 - Airborne Auxiliary Power 49-60 APU Control
DETAILED COMPONENT INFORMATION CAUTION
DATA MEMORY MODULE The data memory module (DMM) interfaces with the APU ECU. The APU ECU manages the communication between the OMS and DMM. The DMM stores information regarding the health monitoring and life usage of the APU. The DMM interfaces with the ECU via a serial link and stores the following information: •
APU serial number and model number
•
APU operating hours and starts
•
APU low cycle fatigue and creep life
•
APU accumulated aborted starts
Do not remove the APU electronic control unit (ECU) and data memory module (DMM) at the same time. Always do the APU BIT check between replacement of each of the units or data may be erased.
At power-up, the ECU reads the DMM and performs the following: •
If the information is the same as the ECU, power-up initialization occurs
•
If the information is different, the ECU record is updated from the DMM
•
Upon completion of an APU cooldown cycle, the DMM is automatically updated from the ECU
The APU data is also stored in the ECU in case of DMM failure. Replacement DMMs are delivered with all zeros stored in their memory. Once recognized, the ECU records in memory that a DMM change has occurred. The new DMM is then loaded with all the totaled APU data that is stored in the ECU. If the ECU is replaced, the DMM retains all the APU data and automatically re-initializes the new ECU with the old accumulated APU data.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
APU
OMS
DMC 1 and DMC 2
Data Memory Module
CS1_CS3_4961_004
ELECTRONIC CONTROL UNIT
LEGEND ARINC 429 RS 485 Discrete
Figure 46: Data Memory Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
DETAILED DESCRIPTION ECU POWER The ECU primary power is supplied by DC ESS BUS 2. The ECU also has a secondary power input that is connected to the aircraft DC EMER BUS through the APU relay controlled by the ECU. The dual input power configuration allows the ECU to select the highest voltage power source, remain powered ON whenever the DC EMER BUS is switched on, and allows the DMC to continuously communicate with the ECU whenever the aircraft is powered. APU Start During normal operation, the ECU is powered on via the DC ESS BUS 2 when the aircraft battery switch is closed. After the APU switch is turned to RUN and the ECU completes successful prestart BIT, the ECU applies power to the APU relay and provides secondary power to the ECU via the DC EMER BUS. APU Shutdown On the ground, once the APU switch is turned to OFF and the APU 60 second cooldown, shutdown, and memory management activities are completed, the ECU de-energizes the relay. The APU ECU power is only supplied by the DC ESS BUS 2. In flight, the relay remains activated at APU shutdown to retain power input redundancy.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
EPC 1 APU Relay
APU Inlet Door Pwr DC EMER BUS
10
APU ECU Sec DC EMER BUS
10
APU ECU
28 VDC Ground
APU RUN
START
Start Command
PULL TURN
APU PANEL
CDC 2-8-1 APU ECU PRI DC ESS BUS 2 SSPC 10A Logic: Always On
CS1_CS3_4961_010
OFF
Figure 47: ECU Power Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
ECU INTERFACE The ECU communicates with the IASC to determine the APU bleed mode required. The ECU positions the IGVs and SCV based on this information.
The groomer mode is entered when the maintenance outlets in either galley 1 or 4 are used. In the groomer mode, the APU is governed to a frequency of 400 Hz.
The ECU receives DC electrical power from DC ESS BUS 2 and the DC EMER BUS. The ECU receives a signal from the fire detection and extinguishing (FIDEX) system when an APU fire is detected. When pressed, the APU FIRE pushbutton annunciator (PBA) also closes the shutoff valve (SOV), independent of the ECU. The ECU controls the APU fuel shutoff valve during normal operation. The ECU provides information to the EICAS and synoptic pages and receives input signals from aircraft systems via an ARINC 429 BUS from the data concentrator unit module cabinets (DMCs). The DMCs provide: •
Electrical load
•
Electrical load shed requests
•
Main engine start requests
•
Engine status
•
Weight-on-wheels (WOW) status
•
Air data
•
APU switch commands
Unless otherwise specified, when duplicate data is available from both DMCs, the ECU selects DMC 1 as the primary data source. The ECU controls and receives position feedback from APU inlet door actuator. The APU can be shut down externally by a switch located in the external service panel or from the switch in the APU compartment.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
LEGEND ADSP AGCU APU CDC DMC ECU EPC FADEC IASC LGSCU
Air Data System Probe Auxiliary Generator control Unit Auxiliary Power Unit Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Control Unit Electronic Power Control Full Authority Digital Engine Control Integrated Air System Controller Landing Gear and Steering Control Unit
EPC 1 APU RELAY
APU ECU
Secondary Power Relay Low
CDC 2-8-1 APU ECU PRI DC ESS BUS 2 SSPC 10A Logic: Always On
APU ECU SEC
Secondary 28 VDC
10 CDC 1
400 Hz on Ground CDC 5
CDC 1-5-16 APU SOV AUTO DC ESS SSPC 3A BUS 1 Logic: Combinational
DC EMER BUS
IASC 1 and IASC 2
APU FUEL SOV DC 3 EMER BUS
LGSCU
APU FIRE PANEL
EXTERNAL SERVICE PANEL
K6
AGCU Ready to Load
APU OVERHEAD SWITCH
Secondary Power Relay High
APU FIRE PBA Generator Load in
APU FEED SOV Close
APU Fuel SOV Command Close 28 VDC Primary Power APU Fuel SOV Command Open
DMC 1 and DMC 2
APU Shutdown
EXTERNAL TOWING SERVICE PANEL
APU FIRE SW
Pos Open
FIDEX CONTROL UNIT
28 VDC Return APU INLET DOOR
Excitation 28 VDC
APU COMPARTMENT SHUTDOWN SWITCH
LEGEND ARINC 429 AFDX
APU Shutdown
LEFT EEC and RIGHT EEC
CS1_CS3_4961_016
ADSP
Feedback
Figure 48: Electronic Control Unit Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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49 - Airborne Auxiliary Power 49-60 APU Control
APU PROTECTIVE SHUTDOWNS The ECU provides protective, automatic shutdowns to prevent APU or aircraft damage. The weight-on-wheels (WOW) signal is used to inhibit some of the auto-shutdown functions when the APU is in flight mode. The ECU automatically shuts down the APU when certain fault conditions are detected. Protective shutdowns are indicated on the engine indication and crew alerting system (EICAS) by an APU SHUTDOWN advisory message with the exception of OVERSPEED, which is a dedicated warning message. In flight, the ECU inhibits some protective shutdown. When a protective shutdown is inhibited in flight, an APU caution message is displayed. The APU can be shut down manually if it is not required. During a protective shutdown, there is no cooling period and the APU shuts down immediately. When a protective shutdown occurs, the electrical and pneumatic loads are removed immediately. The bleed air valve closes and the SCV opens. Power is removed from the APU fuel control unit fuel solenoid and torque motor, and then the APU fuel SOV closes. The APU switch must be manually selected OFF when a protective shutdown occurs.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
Table 6: APU Protective Shutdown CONDITION
AUTO SHUTDOWN ON GROUND
AUTO SHUTDOWN IN FLIGHT
Failed/slow/hot/hung start
YES
YES
RPM underspeed
YES
NO
RPM overspeed/signal loss
YES
YES
Critical ECU fault (i.e. loss of DC power)
YES
YES
Loss of ARINC inputs to the ECU
YES
YES
APU fire
YES
NO
EGT overtemperature/signal loss
YES
NO
APU door position uncommanded/signal loss YES
NO
Excessive oil temperature
YES
NO
Low oil pressure
YES
NO
Oil filter clogged
YES
NO
APU inlet overheat/reverse flow
YES
NO
CS1_CS3_TB49_002
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the APU control system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
CAS MESSAGES Table 10: INFO Messages
Table 7: WARNING Message MESSAGE
LOGIC
APU OVERSPEED
MESSAGE
APU overspeed and no auto-shutdown.
49 APU FAULT - APU INOP
APU inoperative (APU fuel solenoid did not close when commanded fuel solenoid or loss of ARINC 429 input from DMC) or shutdown condition is detected.
49 APU FAULT - APU DEGRADED
Oil temp sensor fault or inlet temp sensor fault or ECU noncritical fault or APU low fuel flow and not inflight or oil pressure switch is lost.
49 APU FAULT - APU REDUND LOSS
APU secondary power relay command open short or APU 28 VDC secondary low or APU 28 VDC primary low and not APU 28 VDC secondary power relay command open short or APU 28 VDC secondary low and APU 28 VDC primary low or APU speed sensor failure and not loss of both speed sensors or EGT1 loss or EGT 2 loss and not EGT 1 and EGT 2 loss or APU ARINC 429 to DMC 1 no data received or APU ARINC DMC 2 no data received and not loss of both ARINC 429 signals.
Table 8: CAUTION Message MESSAGE
LOGIC
APU
APU failure in flight that does not lead to an automatic shutdown.
Table 9: ADVISORY Message MESSAGE
LOGIC
APU SHUTDOWN
Failures detected that cause the APU to shut down (mostly used on ground).
APU FAULT
Loss of redundant or noncritical function for the APU systems.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
TECHNICAL TRAINING MANUAL
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49 - Airborne Auxiliary Power 49-60 APU Control
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM The APU fuel shutoff valve is tested through the onboard maintenance system (OMS). The fuel shutoff valve test is found on the APU test page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-110
CS130_49_L2_APUControl_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-60 APU Control
MAINT MENU
RETURN TO LRU/SYS OPS
APU ECU-FUEL SHUT OFF VALVE TEST
001 /006 APU-FUEL SHUT OFF VALVE TEST PRECONDITIONS 1 - SET `APU RUN SW' TO OFF (APU SHUTDOWN) 2 - VERIFY THAT WEIGHT ON WHEELS IS TRUE
TEST INHIBIT/INTERLOCK CONDITIONS APU SHUTDOWN:
YYYYY
WEIGHT ON WHEELS:
YYYYY
PRESS `CANCEL' BUTTON TO GO TO MAIN MENU
Continue
Write Maintenance Reports
Cancel
CS1_CS3_4931_008
PRESS `CONTINUE' BUTTON TO GO TO NEXT PAGE
Figure 49: OMS Fuel Shutoff Valve Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-111
CS130_49_L2_APUControl_PracticalAspects.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
49-00 APU OPERATION GENERAL DESCRIPTION The auxiliary power unit (APU) is started from the flight deck using the APU switch located on the overhead panel. When the APU switch is momentarily held in the START position, the APU begins the start sequence: •
Built-in test (BIT)
•
APU door open
•
APU feed shutoff valve (SOV) open
When released, the selector automatically returns to RUN position. If the BIT tests are successful, the APU start is initiated. The electronic control unit (ECU) controls the APU start sequence. The ECU continuously monitors the APU start sequence and shuts down the APU if a fault is detected. If a BIT fails or a fault is detected during start, the APU switch is selected to OFF and a restart can be carried out after a two minute cooldown period. When the rotary selector is placed in the RUN position after START, the APU runs automatically, providing air and electrical loads as required. On ground, moving the APU switch to the RUN position without going to START enables the maintenance mode, which opens the APU inlet door and the aircraft APU feed shutoff valve. When the APU is running, selecting the APU switch to OFF initiates the cooldown sequence and the APU shuts down.
NOTE Moving the selector between OFF and RUN position requires a pull-to-turn action. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-112
CS130_49_L2_APUOperation_General.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU Switch Off
Start Sequence or Maintenance Mode
Start Sequence
APU Switch to START, then to RUN BIT Fail
Maintenance Mode
BIT Fail
APU Switch to RUN
BIT Starts BIT Intiated, Inlet Door and Bit Test APU Fuel SOV Open 5 Seconds BIT OK APU Switch to START, Release to RUN Select OFF Restart After 2 Minutes
APU Failed to Start
BIT OK
APU Start Sequence
Shutdown APU by Selecting Switch to OFF
APU Off
CS1_CS3_4961_017
APU Runing
Figure 50: APU Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-113
CS130_49_L2_APUOperation_General.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
CONTROLS AND INDICATIONS FLIGHT DECK CONTROLS APU Fire Panel The APU FIRE PBA is located on the ENGINE and APU FIRE panel on the overhead integrated cockpit control panel (ICCP). The APU FIRE PBA provides 28 VDC to close the APU feed shutoff valve through hardwired switch contacts. APU Panel The APU master switch is located on the left inboard ICCP module of the overhead panel. The switch position is sent to the ECU via an ARINC 429 BUS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-114
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU
APU FIRE
OFF
RUN
START
PULL TURN
BTL
APU PANEL
CS1_CS3_4961_011
ENGINE AND APU FIRE PANEL
Figure 51: APU Flight Deck Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-115
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
EXTERNAL CONTROLS There are two external APU shutdown switches. Operating either switch causes the APU to shut down without a cooldown period. The switch inputs are inhibited when the aircraft is in flight. ELECTRICAL/TOWING SERVICE Panel The APU SHUT OFF toggle switch on the external service panel is a guarded momentary contact switch with a center OFF position. The switch is hardwired to the data concentrator unit module cabinet (DMC). APU Compartment Shutdown Switch The APU compartment shutdown switch is mounted on the forward firewall. The switch is hardwired to the APU ECU.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-116
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
MAINT LTS ON/OFF
NAV LTS
Shutdown Switch
APU SHUT OFF
LAMP TEST
CKPT IN USE EXT PWR
TOWBARLESS ONLY
AVAIL
NO TOW
IN USE
TOW
SERV
BATT
ON TOW PWR
OMS/HMU OMS/ MS/HMU
C CALL
ON
HEADSET
APU COMPARTMENT SHUTDOWN SWITCH
CS1_CS3_4961_013
ELECTRICAL/TOWING SERVICE PANEL
Figure 52: APU External Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-117
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
SYNOPTIC PAGE The APU performance information is displayed on the STATUS synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-118
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
STATUS
AIR IR
DOOR OR
ELEC C
FLT CTR CTRL L
FUEL FUE EL
HYD YD
AVIONIC IONIC
INFO
CB
15 °C 15 °C
TAT SAT
APU RPM Symbol
ENGINE
103 113 21.8
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L) APU
RPM EGT DOOR
100 % 312 °C OPEN
OIL TEMP OIL PRESS OIL QTY
85
APU operation or below redline.
107
APU operation above redline (overspeed).
è½½
APU rpm value invalid.
75 °C NORM NORM APU EGT
TIRE PRESSURE (PSI) Symbol
150 224
150
224
224
Condition
224
Condition
650
APU operating at or below redline.
960
APU operating above redline (overtemperature).
è½½
APU EGT invalid.
BRAKE
00
TEMP
01
01
CS1_CS3_4991_011
00
SYNOPTIC PAGE - STATUS
Figure 53: APU Status Performance Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-119
CS130_49_L2_APUOperation_ControlsIndications.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
•
OPERATION APU START On the APU control panel, pull and turn the APU switch to the START position and release to RUN. •
APU inlet door opens
•
On the STATUS synoptic page:
•
-
APU DOOR indicates open
-
APU RPM increases
-
RPM 91 to 100% (ground) 100% (flight)
-
EGT less than 950°C
-
DOOR OPEN
-
OIL TEMP NORM
-
OIL PRESS NORM
-
OIL QTY FULL
WARNING
On EICAS: -
•
On STATUS synoptic page:
APU IN START status message
On the FUEL synoptic page: -
APU feed SOV open
-
APU flow line green
-
APU fuel filter green
BEFORE OPERATING THE APU, MAKE SURE THAT THE ACCESS DOORS ARE LOCKED IN THE OPEN OR CLOSED POSITION. DO NOT OPEN OR CLOSE THE ACCESS DOORS DURING APU OPERATION. THE APU AIR SUCTION CAN CAUSE THE DOORS TO CLOSE QUICKLY IF THEY ARE NOT LOCKED AND CAUSE INJURY TO PERSONS AND DAMAGE TO THE EQUIPMENT.
NOTE APU IN START is displayed until APU reaches 70% rpm. At 70% rpm, the following message is shown on EICAS: •
APU ON status message
At 100% rpm: •
On EICAS -
APU ON status message
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
CAUTION 1. The APU start from battery shall only be allowed if there are two batteries available and TRU 3 is not available. A battery start is allowed with one battery available and a TRU feeding the DC BUSES when TRU 3 is not available. 2. The APU fire extinguishing system is not effective when the access doors are open. If the APU is operated with the access doors open, make sure that fire fighting equipment is available.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-120
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU RUN
OFF
APU
NOTE
START
OFF
Do not perform more than three starts/start attempts in one hour. A two-minute delay must be observed between start attempts.
PU TU LL RN
RUN
START
PULL TURN
APU PANEL
APU PANEL
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
4000 KG 0 KG
20 °C 2000 KG
20 °C 2000 KG
APU IN START
APU ON
APU ON 0 KG
EICAS
EICAS
EICAS
APU
EGT DOOR
0% 22 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
25 °C NORM FULL
RPM EGT DOOR
SYNOPTIC PAGE – STATUS
SYNOPTIC PAGE – FUEL
70 % 285 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
32 °C NORM NORM
RPM EGT DOOR
100 % 315 °C OPEN
OIL TEMP OIL PRESS OIL QTY
97 °C NORM NORM
SYNOPTIC PAGE – STATUS
SYNOPTIC PAGE – STATUS
70% rpm
100% rpm
0% rpm
CS1_CS3_4961_019
APU RPM
Figure 54: APU Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-121
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
NORMAL APU SHUTDOWN The normal shutdown is carried out using the APU switch in the flight deck. Before shutting down, remove the electrical and pneumatic loads. On the APU control panel, pull and turn the APU switch to the OFF position. After one minute: •
APU stops, RPM decreases to 0%, and EGT decreases towards ambient temperature
•
On the FUEL synoptic page, the APU SOV shows CLOSED, and the APU flow line shows white
•
On the STATUS page DOOR CLOSED is shown in white
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-122
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU OFF
APU
RUN
START
OFF
RUN
LL PU N R TU
START
LL PU N R TU
60 Second Time Delay APU PANEL
APU PANEL
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
4000 KG 0 KG
60 Second Time Delay Ends
20 °C 2000 KG
20 °C 2000 KG
APU ON EICAS
0 KG
EICAS APU
EGT DOOR
100 % 315 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
97 °C NORM NORM
SYNOPTIC PAGE – STATUS
RPM EGT DOOR
0% 125 °C CLOSED
OIL TEMP OIL PRESS OIL QTY
32 °C NORM FULL
SYNOPTIC PAGE – STATUS
100% RPM
SYNOPTIC PAGE – FUEL 0% RPM
CS1_CS3_4961_020
APU RPM
Figure 55: APU Normal Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-123
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
EMERGENCY APU SHUTDOWN The emergency shutdown is carried out from the external shut down switches or the APU FIRE PBA in the flight deck. Pressing any emergency shutdown switch: •
Immediately APU stops with no cooldown period, RPM decreases to 0%, and EGT decreases toward ambient temperature
•
On FUEL synoptic page, APU SOV shows CLOSED and APU flow line shows white
•
On STATUS page DOOR CLOSED is shown in white
On the APU control panel, pull and turn the APU switch to the OFF position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-124
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
APU
APU
APU OFF
RUN
APU
START
OFF
FIRE
RUN
START
FIRE PULL TURN
BTL
PULL TURN
BTL
AVAIL
AVAIL
APU PANEL
APU PANEL ENGINE AND APU FIRE PANEL
ENGINE AND APU FIRE PANEL
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
4000 KG 0 KG
No Cooldown Period
20 °C 2000 KG
20 °C 2000 KG
APU ON EICAS
0 KG
EICAS APU
EGT DOOR
100 % 315 °C OPEN
APU OIL TEMP OIL PRESS OIL QTY
97 °C NORM NORM
SYNOPTIC PAGE – STATUS
RPM EGT DOOR
0% 125 °C CLOSED
OIL TEMP OIL PRESS OIL QTY
32 °C NORM FULL
SYNOPTIC PAGE – STATUS
100% RPM
SYNOPTIC PAGE – FUEL 0% RPM
CS1_CS3_4961_021
APU RPM
Figure 56: APU Emergency Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
49-125
CS130_49_L2_APUOperation_Operation.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
MONITORING AND TEST The following page provides the crew alerting system (CAS) messages for the APU operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-126
CS130_49_L2_APUOperation_MonitoringTest.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
CAS MESSAGES Table 11: STATUS Messages MESSAGE
LOGIC
APU ON
APU is running.
APU IN START
APU performing startup sequence.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-127
CS130_49_L2_APUOperation_MonitoringTest.fm
49 - Airborne Auxiliary Power 49-00 APU Operation
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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49-128
CS130_49_ILB.fm
ATA 71 - Power Plant
BD-500-1A10 BD-500-1A11
71 - Engine
Table of Contents 71-10 Nacelles........................................................................71-2 General Description ..........................................................71-2 71-60 Inlet Cowl .....................................................................71-4 General Description ..........................................................71-4 71-11 Fan Cowls ....................................................................71-6 General Description ..........................................................71-6 Practical Aspects ..............................................................71-8 78-30 Thrust Reverser Doors ...............................................71-12 General Description ........................................................71-12 Component Location ......................................................71-24 Component Information ..................................................71-26 Detailed Description ........................................................71-30 Practical Aspects ............................................................71-32 78-11 Turbine Exhaust System ............................................71-38 General Description ........................................................71-38 71-40 Engine Mounts and Pylon...........................................71-40 General Description ........................................................71-40 Component Location ......................................................71-40 Detailed Component Information ....................................71-42 Practical Aspects ............................................................71-46 71-70 Engine Drains .............................................................71-54 General Description ........................................................71-54 71-00 Engine Storage and Preservation..............................71-56 General Description ........................................................71-56
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-i
CS130_M5_71_PowerPlantTOC.fm
71 - Engine
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7: Figure 8: Figure 9: Figure 10: Figure 11: Figure 12:
Nacelle Systems....................................................71-3 Inlet Cowl...............................................................71-5 Fan Cowls .............................................................71-7 Fan Cowl Opening and Closing.............................71-9 Fan Cowl Latch and Keeper Adjustment.............71-11 Thrust Reverser Doors ........................................71-13 Thrust Reverser Fixed Structure .........................71-15 Thrust Reverser Door Latches ............................71-17 Bumpers and Bifurcation Latch System ..............71-19 Power Door Operating System............................71-21 Power Door Operating System Control ...............71-23 Power Door Operating System Component Location ...........................................71-25 Figure 13: Power Door Operating System Powerpack...........................................................71-27 Figure 14: Power Door Operating System Locking Actuator and Hold Open Rods ...............71-29 Figure 15: Power Door Operating System Schematic..........71-31 Figure 16: Thrust Reverser Door Bifurcation Latch System and Latch Operation .....................71-33 Figure 17: Thrust Reverser Door Opening and Closing .......71-35 Figure 18: Thrust Reverser Door Bifurcation Latch System Adjustment....................................71-37 Figure 19: Turbine Exhaust System .....................................71-39 Figure 20: Engine Mounts and Pylon ...................................71-41 Figure 21: Forward Engine Mount ........................................71-43 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 22: Aft Engine Mount and Thrust Links..................... 71-45 Figure 23: Engine Removal Bootstrap Equipment............... 71-47 Figure 24: Engine Cradle, and Transport Stand .................. 71-49 Figure 25: Electrical Harness Disconnects .......................... 71-51 Figure 26: Hydraulic and Fuel Connections......................... 71-53 Figure 27: Engine Drains ..................................................... 71-55 Figure 28: Engine Storage and Preservation....................... 71-57 Figure 29: Engine Preservation Chart.................................. 71-59
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-ii
CS130_M5_71_PowerPlantLOF.fm
71 - Engine
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-iii
CS130_M5_71_PowerPlantLOF.fm
POWER PLANT - CHAPTER BREAKDOWN POWER PLANT CHAPTER BREAKDOWN
Nacelle Systems
1
Power Door Operating System
2
Engine Mounts and Pylon
3
Engine Storage - Preservation
4
71 - Power Plant 71-10 Nacelles
71-10 NACELLES GENERAL DESCRIPTION The engine cowlings minimize airflow turbulence around the engine. They provide an aerodynamically smooth surface around the engine and protect the engine mounted components. The engine inlet cowl assembly directs the airflow to the engine fan. The inlet cowl is acoustically treated for noise reduction. The fan cowls, installed between the inlet cowl and the thrust reverser doors, are attached to the pylon fan cowl support beam. Each inboard fan cowl has a strake installed to decrease the turbulence caused by the airflow between the fan cowl and the fuselage. The thrust reverser doors enclose the core engine and contain the necessary components for the thrust reverser system. The thrust reverser doors attach to the hinge beam on the pylon. Both the fan cowls and thrust reverser doors can be opened to provide access to the engine. The turbine exhaust system controls and directs the exhaust gases.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-2
CS130_71_L2_Nacelles_General.fm
71 - Power Plant 71-10 Nacelles
Strake
Pylon
Turbine Exhaust System
Fan Cowl
Thrust Reverser Door CS1_CS3_7100_015
Inlet Cowl
Figure 1: Nacelle Systems Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-3
CS130_71_L2_Nacelles_General.fm
71 - Power Plant 71-60 Inlet Cowl
71-60 INLET COWL GENERAL DESCRIPTION The inlet cowl provides smooth undisturbed airflow to the fan in all operating conditions, and reduced noise. The inner barrel of the inlet cowl is made of titanium and composite with 360° of interior acoustic treatment. Lightning strike protection is provided by a copper screen layer impregnated into the outer barrel. The outer barrel is a two-piece composite structure extending from the inlet lip skin to the leading edge of the fan cowl, providing a smooth surface for airflow across the nacelle exterior. Two hoist points are provided for the removal of the inlet cowl. A NACA scoop at the 1 o’clock position is used for fan compartment cooling. A panel, located at the 11 o’clock position, provides access to the P2T2 probe, sense line, and wiring harness. The inlet cowl is anti-iced by bleed air. The air is exhausted out of the bottom of the inlet cowl through a series of slots. An access panel for the cowl anti-ice (CAI) duct is located at the 5 o’clock position. Drain holes in the core of the inner barrel and at the bottom of the outer barrel allow fluids to drain. The inlet cowl is bolted to the front flange of the engine at the attachment ring.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-4
CS130_71_L2_InletCowl_General.fm
71 - Power Plant 71-60 Inlet Cowl
P2T2 Probe
Hoist Points
NACA Scoop
Inner Barrel Sound Proofing
P2T2 Access Panel
Outer Barrel
Attachment Ring Inlet Cowl Sling
Anti-ice Air Exhaust Louvers
Cowl Anti-Ice Duct Access Panel
Lip
CS1_CS3_7111_002
Drain Holes
Figure 2: Inlet Cowl Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-5
CS130_71_L2_InletCowl_General.fm
71 - Power Plant 71-11 Fan Cowls
71-11 FAN COWLS GENERAL DESCRIPTION The fan cowls cover the fan case, providing a protective enclosure for engine mounted components and a smooth aerodynamic outer surface for the nacelle. They are hinged at the top and latched at the bottom centerline with three flush mounted latches. The fan cowl doors are manufactured using carbon fiber reinforced polymer (CFRP), strengthened with radial stiffeners and longerons. Copper mesh is embedded in the fan cowl laminate for lightning strike protection. The removable strakes are installed on the inner cowls only. A pressure-relief door, located on the outer fan cowls, relieves high-pressure in the event of a duct failure. During maintenance, hold open rods (HORs) provide a means to support the cowls in the open position. Drain holes, located near the lower centerline, ensure fluid drainage. Three hoist points of two screw holes are on each panel. They are used to attach the fan cowl door assembly to the ground handling sling. The ground handling sling is used to lift the cowl during removal or installation. Wear strips ensures aerodynamic smoothness is maintained between the cowls. Three hinges attach the cowls to the pylon. These hinges are secured by pins held in place by retaining clips. The pins facilitate removal of the fan cowl from the pylon. Two HORs on each panel are used to hold the cowls in the open position. Quick-release fittings secure one end of the rods to the fan case when the cowls are opened. The other end of the rods are attached to a pivot bracket assembly. Stow brackets and clips on the cowl provide a means to secure the rods when the cowl is closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-6
CS130_71_L2_FanCowls_General.fm
71 - Power Plant 71-11 Fan Cowls
Stiffeners
Quick-Release Fitting Hinges Hold Open Rod Clip
Pressure-Relief Door
Strake
Wear Strips
Hoist Attachment Points
Latches
HOLD OPEN ROD
CS1_CS3_7111_004
Drain Holes
Figure 3: Fan Cowls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-7
CS130_71_L2_FanCowls_General.fm
71 - Power Plant 71-11 Fan Cowls
PRACTICAL ASPECTS
•
Always refer to the latest Aircraft Maintenance Publication (AMP) before carrying out any procedures on the aircraft.
Put the forward HOR on to the HOR clip. Secure the HOR with the snap spring of the clip
•
Carefully lower the fan cowl and push the fan cowl door together at the bottom centerline. Make sure the alignment pins enter the holes adjacent to the latches
•
Engage the hooks on all three latches with the eye-bolts on the three latch keepers in the sequence L1, L3, L2
•
Close the three latch handles until they are flush with the door surface and locked in position. When the latches are properly closed, the red marks on the latches do not show
OPEN THE FAN COWL To open the fan cowl: •
Push the fan cowl door latch triggers to release the three latch handles. Pull down each handle to open the three latch handles
•
Move the latches away from the three latch keepers in the sequence L1, L3, and L2
•
Manually lift and hold a fan cowl door at the lower edge
•
Remove the forward hold open rod (HOR) from the fan cowl and install it in the forward HOR bracket on the fan case. Secure it with the quick-release pin on the end of the HOR
•
Remove the aft HOR from the fan case and secure it to the power door operating system (PDOS) bracket on the fan case. Secure it with the quick-release pin on the end of the HOR
•
Slowly lower the fan cowl door until the HORs hold the weight of the door
CLOSE THE FAN COWL To close the fan cowl: •
Manually lift and hold the fan cowl door at the lower edge so that the weight is not on the HORs
•
Remove the aft HOR from the PDOS bracket on the fan case and install it on the fan cowl. Secure the HOR with the quick-release pin
•
Remove the forward HOR from the forward HOR bracket on the fan case and install it on the fan cowl. Secure the HOR with the quick-release pin
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
WARNING 1. TWO OR MORE PEOPLE ARE REQUIRED TO OPEN THE FAN COWL DOOR. THE FAN COWL DOOR WEIGHS MORE THAN 45.36 KG (100 LB). IF THE DOOR FALLS, INJURY AND/OR DAMAGE TO THE ENGINE OR PERSON(S) CAN OCCUR. 2. DO NOT KEEP OPEN A FAN COWL DOOR WHEN THE WIND SPEED IS OVER 96 KM/H (60 MPH). IF THE WIND MOVES THE FAN COWL DOOR, INJURY AND/OR DAMAGE TO THE ENGINE OR PERSON(S) CAN OCCUR. 3. CAUTION IS REQUIRED IF YOU OPEN A FAN COWL DOOR WHEN THE WIND SPEED IS 74 KM/H (46 MPH) OR MORE. IF THE WIND MOVES THE FAN COWL DOOR, INJURY AND/OR DAMAGE TO THE ENGINE OR PERSON(S) CAN OCCUR.
CAUTION Lifting the fan cowl door more than 55° vertical can cause fan cowl door damage or pylon damage.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-8
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 71-11 Fan Cowls
A
LATCH SEQUENCE OPENING: L1, L3, L2 CLOSING: L1, L3, L2
FAN COWL OPEN (VIEW LOOKING FORWARD)
CS1_CS3_7111_010
C05210065-001
A
Figure 4: Fan Cowl Opening and Closing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-9
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 71-11 Fan Cowls
FAN COWL LATCH AND KEEPER ADJUSTMENT The fan cowl latches require adjustment anytime the fan cowl does not close properly and the latches are safely engaged. The latches engage adjustable latch keepers on the opposite cowl. If the fan cowl latches require adjustment, the keepers are adjusted by inserting an allen key into the star wheel adjustment nut. The latches are adjusted to provide a closing force of 9.1 to 10.9 kgf (20 to 24 lbf). Shim plates on latches and keepers are used to adjust the cowl gap. Make sure the gap between the left and right fan cowls is 0.51 - 5.08 mm (0.020 - 0.200 in.).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-10
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 71-11 Fan Cowls
Shim Plates
Fan Cowl Latch Keepers B
C
A A Fan Cowl Latches
Star Wheel Adjuster
Keeper Hook
Latch
Star Wheel Adjuster
Allen Key
C
B
CS1_CS3_7111_003
Latch Handle
Figure 5: Fan Cowl Latch and Keeper Adjustment Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-11
CS130_71_L2_FanCowls_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
78-30 THRUST REVERSER DOORS GENERAL DESCRIPTION The power plant is equipped with a cascade type thrust reverser (T/R) assembly consisting of left and right T/R doors. The left and right T/R doors are installed on a hinge beam on the pylon. The hinge beam has four attachment points. The T/R doors are latched together at the bottom of the engine and form an enclosure for the core engine. The T/R doors can be opened using the power door operating system (PDOS), or manually using a hydraulic pump. When opened, the engine core can be accessed. Each T/R door consists of two fixed structures that direct fan air aft. The inner structure forms an aerodynamic surface from the fan case exit to the aft nozzle. The outer structure contains the T/R translating sleeve, cascades, and blocker doors that provide reverse thrust. The inner and outer fixed structures are connected by the torque box. The torque box is the attachment point for the T/R components and the PDOS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-12
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Hinge Beam
Blocker Door
Cascade Segments
Torque Box
Inner Fixed Structure Latch Beam
CS1_CS3_7110_001
Translating Sleeve
Figure 6: Thrust Reverser Doors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-13
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
THRUST REVERSER FIXED STRUCTURE The thrust reverser fixed structure covers the engine core. It provides core compartment ventilation, and forms the fan duct inner aerodynamic surface from the fan case exit to the exhaust nozzle. The thrust reverser attaches to the pylon at the hinge beam. An air/oil cooler (AOC) inlet and exhaust is located on the left side of the fixed structure. A pressure-relief door, located behind the AOC, opens in the event of a burst duct in the bleed air system. The pressure relief door is located on the left side of the inner fixed structure behind the air/oil cooler (AOC) exhaust, at the 9 o'clock position. The latches release at 3.5 psi. The door is manually closed if it has opened. The right side of the fixed structure has an oil tank service door and a cutout for the active clearance control (ACC) ducting. A precooler exhaust door is located behind the oil tank service door. Forward and aft bumpers, located at the top and bottom of the structure, are used to align the doors as they are closing. A locking latch is used to ensure the doors remain closed in the aligned position. Latches, installed on the latch beam, are used to secure the T/R door in the closed position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-14
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
AOC Inlet and Exhaust
Hinge Beam
Fwd Upper Bumper
ACC Duct
Aft Upper Bumper
Oil Tank Service Door Precooler Exhaust Door
Locking Latch
Aft Upper Bumper
Fwd Lower Bumper Pressure-Relief Door Latch Beam Aft Latch
Aft Lower Bumper
Aft Lower Bumper
LEFT THRUST REVERSER FIXED STRUCTURE
RIGHT THRUST REVERSER FIXED STRUCTURE
CS1_CS3_7111_007
LEGEND ACC AOC
Aft Latch
Active Clearance Control Air/Oil Cooler
Figure 7: Thrust Reverser Fixed Structure Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-15
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
THRUST REVERSER DOOR LATCHES Latches are quick-release mechanisms integral to the latch beam that provide quick access to the engine core. The latches provide resistance to loads that might otherwise cause the T/R to disengage or open during the flight cycle. Latches no. 1, no. 2, no. 3, and no. 4 are installed on the latch beam. Latch no. 5 is located on the lower aft section of the inner fixed structure. The latches are takeup latches that pull the two T/R halves together when securing. Latches no. 2 and no. 3 are located inside the latch access door. The latch access door cannot be closed if latch 2 and latch 3 are not properly secured. A closure assist assembly consists of a turnbuckle, which is used to draw the door together before engaging the latches. An eye on one end of the turnbuckle allows it to pivot in order for the pin to engage the opposite cowling. When engaged, the turnbuckle is turned manually to draw the two doors together. The closure assist assembly is stowed after use.
CAUTION 1. Do not turn the body of the closure assist assembly to make it shorter than the minimum length. The closure assist assembly can be damaged if the minimum length is decreased. 2. Adjust the closure assist assembly to the minimum length before placing the retention clip of the stow bracket assembly. The closure assist assembly can move out of position if this step is not followed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-16
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Closure Assist Assembly Stowed Position
Latch No. 2
A
Latch No. 3
Latch No. 5
Latch No. 1 Latch No. 4
Latch Access Door
CS1_CS3_7111_009
A
Figure 8: Thrust Reverser Door Latches Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-17
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Bumpers and Bifurcation Latch System To maintain a correct fit of the ducts to the engine, the inner fixed structure (IFS) has upper and lower bumpers to assist alignment as the two thrust reverser doors are closed. The upper bumper includes a compression strut to compensate for the larger gap between the two cowls. The bifurcation latch system (BLS) is similar to the bumpers but has a locking feature to keep it engaged. The handle is accessed through the latch access door. The BLS is used to limit the deflection of the IFS, and maintain its structural integrity in the event of a burst air duct. When engaged, the bifurcation latch uses a rod and cam mechanism to insert a locking pin through the latch. As the thrust reverser (T/R) doors close, a pin on the left side is inserted into a receiver located on the right side. The pin mechanism is locked together using a rod. The rod is inserted through the pin by turning, and pushing the BLS lever up. The bifurcation latch lever is painted red, providing a visual indication of the position. If the bifurcation latch lever is not locked, the latch access panel cannot be closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-18
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Compression Bumper
Bumpers Latch Fitting
Adjustable Pin
Upper Bumpers
Compression Strut
Rod
Latch Beam Attach
Bifurcation Latch System Handle
CS1_CS3_7100_004
Lower Bumpers
Figure 9: Bumpers and Bifurcation Latch System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-19
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
POWER DOOR OPERATING SYSTEM The thrust reverser (T/R) power door operating system (PDOS) is used to open the T/R doors and support them in the open position. Each T/R door is opened independently of the other. The PDOS powerpack consists of a hydraulic gear pump driven by an AC motor, and a reservoir capable of supplying fluid to two locking actuators. The PDOS is operated by switches located on each side of the fan case. Each T/R door is raised and lowered using a locking actuator connected between the engine fan case and the T/R door torque box. The locking actuator has an internal locking mechanism that is engaged when the actuator is fully extended. A hold open rod (HOR) provides additional support for the T/R door. The locking actuator takes 100% of the load under normal operation, with the HOR as a failsafe mechanism. Each locking actuator is connected to the hydraulic powerpack using a hydraulic line with a quick-disconnect fitting. If the powerpack fails, the quick-disconnect fitting can be removed from the locking actuator and a hand pump connected in its place.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-20
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Locking Actuator
Powerpack
Hold Open Rod Power Door Operating System Control Switch
CS1 CS3 7100 013 CS1_CS3_7100_013
Hydraulic Line
Figure 10: Power Door Operating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-21
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Power Door Operating System Control The PDOS can be operated when the aircraft is on the ground and the engine is shut down. The left engine powerpack motor is powered by 115 VAC from AC BUS 1. The PDOS operates when the thrust reverser (T/R) door switch is pressed. Only one door can be opened at a time. The switch supplies 28 VDC from DC BUS 1 to energize solenoids within the powerpack. When the solenoid is energized, the powerpack supplies 2300 psi hydraulic pressure to the locking actuator. The right engine PDOS motor is powered from AC BUS 2, and the control power is supplied from DC BUS 2. When the actuator fully extends, a pressure-relief valve (PRV) in the actuator opens to ensure the maximum system pressure is not exceeded. The T/R door switch is held for 5 seconds after full actuator travel, then released. The PRV closes, and the internal lock engages and holds the door open. To close the door, the T/R door switch is held. The actuator extends slightly and the internal lock releases. When the lock is released, the switch is released. The actuator retracts due to the weight of the T/R door. An internal restrictor limits the rate of closing. If no power is available, a hand pump can be connected to the actuator quick-disconnect fitting to open the door.
CAUTION 1. The inboard thrust reverser doors can contact the slats if the slats are in the extended position. 2. The actuators can be opened manually using a hand pump. The introduction of additional fluid in the system may cause the powerpack reservoir to overflow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-22
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
DC BUS 1
Left Thrust Reverser Door Switch
Right Thrust Reverser Door Switch
POWERPACK
Left Thrust Reverser Door Actuator
Right Thrust Reverser Door Actuator
Hand Pump NOTE LEGEND
Left engine shown. Right engine similar.
Hydraulic Pressure
CS1_CS3_7100_016
AC BUS 1
Figure 11: Power Door Operating System Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-23
CS130_71_L2_TRD_General.fm
71 - Power Plant 78-30 Thrust Reverser Doors
COMPONENT LOCATION The power door operating system (PDOS) consists of the following components: •
Powerpack
•
Locking actuators
•
Control switches
•
Hold open rods (HORs)
POWERPACK The powerpack is located on the fan case at the 10 o’clock position. LOCKING ACTUATORS The locking actuators are located on the fan case at the 3 o’clock and 9 o’clock positions. CONTROL SWITCHES The control switches are located on the fan case at the 4 o’clock and 8 o’clock positions. HOLD OPEN RODS The hold open rods (HORs) are installed on each thrust reverser (T/R) door torque box.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-24
CS130_71_L2_TRD_ComponentLocation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Powerpack
Hold Open Rod Power Door Operating System Control Switch
CS1_CS3_7100_006
Locking Actuator
Figure 12: Power Door Operating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-25
CS130_71_L2_TRD_ComponentLocation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
COMPONENT INFORMATION POWERPACK The self-contained power door operating system (PDOS) is powered by an electrohydraulic powerpack producing 2300 psi. The powerpack houses an AC motor and a gear pump assembly, which provides hydraulic pressure to two actuators during system operation. The powerpack includes a reservoir, solenoid valves, filter, and a filler cap/strainer. A drain/vent port on the lower half of the powerpack provides venting and overflow drainage. The system is serviced with engine oil. A filler cap on the reservoir lid is used to replenish the oil. A dipstick is provided under the filler cap for checking the oil level. A strainer prevents any contaminants from entering the reservoir. Oil levels are marked on the dipstick as follows: •
L - Low 0.74 L
•
N - Normal 0.82 L
•
H - High 0.89 L
CAUTION Before servicing the PDOS powerpack, the thrust reverser (T/R) doors should be fully closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-26
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Filler Cap/Strainer and Dipstick
Solenoid Valves Drain/Vent Port
CS1_CS3_7100_007
AC Motor
Figure 13: Power Door Operating System Powerpack Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-27
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
LOCKING ACTUATORS Two locking actuators are used independently to open each thrust reverser (T/R) door. Each actuator has a flow control valve (FCV), quick-disconnect connector, pressure-relief valve (PRV), mechanical lock, and a manual bleed feature. The actuators internal flow control valve provides the system with the means to control the closure rate of the cowl when the lock is released. Both actuators have quick-disconnect fittings to accommodate a hand pump in the event of a powerpack failure. When opened, an automatic mechanical lock inside the actuator provides positive retention of the C-duct. A manual bleed port on the actuator is used to expel air from the system. HOLD OPEN RODS Each T/R door has a telescopic hold open rod (HOR) which is attached between the T/R door and the fan case to help hold the T/R doors open. Lock collars are used to extend and retract the HOR. The HOR is used in conjunction with the locking actuator to hold the T/R door open, and provide a safe environment during maintenance. When not in use, the HORs are stored on the T/R door torque box.
WARNING 1. MAKE SURE THAT THE HOLD OPEN RODS (HORS) ARE INSTALLED. FAILURE OF THE POWER DOOR OPERATING SYSTEM (PDOS) CAN CAUSE THE THRUST REVERSER (T/R) TO CLOSE IF THE HORS ARE NOT INSTALLED. THIS CAN CAUSE INJURY TO PERSONNEL OR DAMAGE TO EQUIPMENT.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-28
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Mechanical Lock Pin Rod End
Manual Bleed Port
Cylinder
A
Bearing Quick-Disconnect Fitting A LOCKING ACTUATOR (RETRACTED)
Lock/Release Collar
Unlock
HYDRAULIC HAND PUMP
Lock HOLD OPEN ROD
CS1_CS3_7100_008
Extend/Retract Collar
Figure 14: Power Door Operating System Locking Actuator and Hold Open Rods Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-29
CS130_71_L2_TRD_ComponentInformation.fm
71 - Power Plant 78-30 Thrust Reverser Doors
DETAILED DESCRIPTION The power door operating system (PDOS) is powered when the aircraft is on the ground and the engine is shut down. For the left engine, the following logic is supplied to the control and distribution cabinet 1 (CDC 1). •
Weight-on-wheels (WOW) signal from the landing gear and steering control units (LGSCUs)
•
Engine shutdown signal from the electronic engine control (EEC)
When the thrust reverser (T/R) door switch is pressed, the solenoid valve is energized, allowing hydraulic fluid to be supplied to the hydraulic pump. When solenoid is energized, logic turns on the L (R) PDOS Motor SSPC and the motor drives the hydraulic pump. The thrust reverser door hydraulic actuator extends to open the door. If the motor should overheat, a thermal switch shuts the motor off.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-30
CS130_71_L3_TRD_DetailedDescription.fm
71 - Power Plant 78-30 Thrust Reverser Doors
ELECTRONIC ENGINE CONTROL
LEFT DOOR REVERSER DOOR SWITCH
DMC 1 AND DMC 2
PDOS POWERPACK
SOL
LGSCU 1 AND LGSCU 2
CDC 1
WOW L ENG OFF
to Left Thrust Reverser Door Actuator
CDC 1-11-12 L PDOS Ctrl DC SSPC 3A BUS 1
Hydraulic Reservoir
Hydraulic Pump to Right Thrust Reverser Door Actuator
Logic: Combinational
L ENG OFF L or R Solenoid Energized
CDC 1-17-4-5-6 L PDOS Motor AC BUS 1 SSPC 7.5A
SOL
WOW
Motor
Logic: Combinational Thermal Switch
LEGEND
SOL
Hydraulic Pressure Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Landing Gear and Steering Control Unit Power Door Operating System Weight-on-Wheels AFDX ARINC 429 Solenoid
RIGHT DOOR REVERSER DOOR SWITCH
NOTE - Left engine shown, right engine similar. - Left thrust reverser door shown operating.
CS1_CS3_7100_009
CDC DMC LGSCU PDOS WOW
Figure 15: Power Door Operating System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-31
CS130_71_L3_TRD_DetailedDescription.fm
71 - Power Plant 78-30 Thrust Reverser Doors
PRACTICAL ASPECTS THRUST REVERSER DOOR BIFURCATION LATCH SYSTEM AND LATCH OPERATION Open the Bifurication Latch System 1. Open the latch access door.
CAUTION The closure assist assembly must be at its minimum length before being put into the retention clip of the stow bracket assembly. If this is not done, the closure assist assembly can move out of position.
2. Turn the bifurication latch system (BLS) latch handle 90° clockwise.
Close the Thrust Reverser Door Latches
3. Pull the BLS latch handle down until it comes to a stop.
1. Make sure that the T/R doors are fully closed.
4. Turn the BLS latch handle 90° counterclockwise.
2. Adjust the closure assist assembly so that it can engage the hook on the fixed panel support on the other T/R half.
Open the Thrust Reverser Door Latches 1. Remove the closure assist assembly from the stow bracket assembly and adjust it to its minimum length. 2. Engage the closure assist assembly in the hook on the fixed panel support on the other thrust reverser (T/R) door.
3. Turn the body of the closure assist assembly with a wrench to pull the two T/R halves closer together. 4. Close the T/R door latches in the sequence that follows: -
LATCH No. 1
3. Turn the body of the closure assist assembly with a wrench to pull the two T/R doors together and relieve the tension on the T/R latches.
-
LATCH No. 2
-
LATCH No. 3
4. Open the T/R door latches in the sequence that follows.
-
LATCH No. 4
-
LATCH No. 5
-
LATCH No. 5
-
LATCH No. 4
-
LATCH No. 3
-
LATCH No. 2
-
LATCH No. 1
5. Put the closure assist assembly into its stowed position Close the Bifurication Latch System 1. Turn the BLS latch handle 90°clockwise.
5. Put the closure assist assembly into its stowed position. Ensure the closure assist assembly is at its minimum length.
2. Push the BLS latch handle up until it comes to a stop. 3. Turn the BLS latch handle 90° counterclockwise. 4. Close the latch access door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-32
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Latch No. 3
Latch Access Door
Latch No. 5
Latch No. 4
Latch No. 2
Closure Assist Assembly Stowed Position
Closure Assist Assembly Engaged Position
Bifurcation Latch System Handle
CS1_CS3_7100_017
Latch No. 1
Figure 16: Thrust Reverser Door Bifurcation Latch System and Latch Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-33
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
THRUST REVERSER DOOR OPENING AND CLOSING Open the Thrust Reverse Door
Close the Thrust Reverser Door
1. Push PDOS operating switch until the locking actuator is fully extended.
1. Release the HOR from the HOR open fitting on the fan case and install the HOR in the HOR stow bracket.
1. Hold the switch for a minimum of 5 seconds after the locking actuator is fully extended and the powerpack pressure-relief valve (PRV) opens.
2. Push the PDOS operating switch until the locking actuator is fully extended.
2. Release the switch and the locking actuator retracts in a controlled descent until the mechanical lock is engaged. 3. Release the top end of the hold open rod (HOR) from the HOR stow bracket and install it in the HOR open fitting on the fan case.
WARNING 1. DO NOT MOVE BETWEEN THE ENGINE AND THE OPEN THRUST REVERSER (T/R) UNTIL THE THE HOLD OPEN ROD (HOR) IS INSTALLED. THE T/R IS HEAVY AND CAN CLOSE QUICKLY IF THE PDOS FAILS. 2. TO PREVENT INJURY TO PERSONNEL OR DAMAGE TO EQUIPMENT, CAUTION IS REQUIRED WHEN OPENING THE T/R IN HIGH WIND.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
3. Hold the switch for a minimum of 5 seconds after the locking actuator is fully extended and the powerpack PRV opens. 4. Release the switch. The powerpack PRV closes and the locking actuator mechanical lock disengages. 5. Make sure that the thrust reverser outer V-groove blade engages with the outer V-groove on the engine during closing. 6. Make sure that the alignment pins correctly engage the receptacles before the thrust reverser doors are fully closed.
CAUTION Make sure that the T/R half area is clear of tools and equipment before closing the door. Failure to do so can cause damage to the T/R half and to the engine.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-34
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
Hold Open Rod Stow Bracket
Hold Open Rod Fitting
CS1_CS3_7100_018
Hold Open Rod
Control Switch
Figure 17: Thrust Reverser Door Opening and Closing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-35
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
BIFURICATION LATCH SYSTEM ADJUSTMENT When engaged, the bifurcation latch uses a rod and cam mechanism to insert a locking pin through the latch. As the inner fixed structure closes, a pin on the left side is inserted into a receiver located on the right side. The pin and latch mechanism are locked together using a rod. The rod is inserted through the pin by turning, and pushing the bifurcation latch system (BLS) lever up. The bifurcation latch is locked in place when the lever is aligned with the fixed structure by pushing the lever in. The pin latch mechanism is adjustable up to 0.635 cm (0.250 in.) in either direction.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-36
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-30 Thrust Reverser Doors
BLS Left Hand Latch Fitting
Right Hand Latch Fitting
0.635 cm (0.250 in.) Retracted
0.635 cm (0.250 in.) Extended Nominal PIN LATCH ADJUSTABILITY
Adjustable Rod Pin Latch Receptacle
BLS Handle
Adjusting Nut
1
LOWER BLS IN OPEN POSITION
LOWER BLS IN CLOSED (LOCKED) POSITION
NOTE 1
To adjust the pin latch a distance of 0.635 cm (0.250 in.).
CS1_CS3_7100_005
Rod (engaged)
Figure 18: Thrust Reverser Door Bifurcation Latch System Adjustment Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-37
CS130_71_L2_TRD_PracticalAspects.fm
71 - Power Plant 78-11 Turbine Exhaust System
78-11 TURBINE EXHAUST SYSTEM GENERAL DESCRIPTION The turbine exhaust system helps control the flow of the exhaust gases. It consists of an exhaust nozzle and centerbody. The centerbody is made up of the forward centerbody and aft centerbody. The exhaust nozzle and centerbody are bolted to the aft engine flange. Crossflow blockers, mounted on the exhaust nozzle at the 2 o’clock and 10 o'clock positions, enhance performance by reducing secondary air flow disturbances across the pylon area. Fireseal fingers, also known as turkey feather seals are mounted on the top of the exhaust nozzle to prevent flames from exiting or entering the core compartment area. During installation, two alignment pins on the forward centerbody facilitate proper clocking of the nozzle. Drain holes help remove any fluid accumulation in the turbine exhaust.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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71-38
CS130_71_L2_TES_General.fm
71 - Power Plant 78-11 Turbine Exhaust System
Exhaust Nozzle Fireseal Fingers
Aft Centerbody
Fwd Centerbody
Crossflow Blocker
Alignment Pins
CS1_CS3_7161_001
Drain Holes
Figure 19: Turbine Exhaust System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-39
CS130_71_L2_TES_General.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
71-40 ENGINE MOUNTS AND PYLON GENERAL DESCRIPTION The engine mounts attach the engine to the aircraft pylon. The nacelles do not carry any significant structural loads. Thrust loads are transferred through the engine mounts. Mounts are disconnected to remove the engine. The function of the engine mount system is to transmit engine and nacelle loads to the pylon. The pylon then transfers these loads to the wing. The main sources of engine mount loads are engine thrust, aerodynamic pressures, and inertia loads from maneuvers and gusts. Thrust links transfer thrust loads to the aft mount and prevent any bending of the engine core. All mounts are designed to be failsafe. Excess clearances in holes along secondary load paths allow the secondary paths to remain inactive, while the primary load paths are intact.
COMPONENT LOCATION Engine mounts located on the fan and turbine exhaust cases are used to attach the engine to the aircraft and pylon. The engine mount system consists of the following components: •
Forward engine mount
•
Aft engine mount
•
Thrust links
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-40
CS130_71_L2_EMP_General.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
A
B
A FWD ENGINE MOUNT
Right Thrust Link
B AFT ENGINE MOUNT
CS1_CS3_7120_001
Left Thrust Link
Figure 20: Engine Mounts and Pylon Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-41
CS130_71_L2_EMP_General.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
DETAILED COMPONENT INFORMATION FORWARD ENGINE MOUNT The forward engine mount is attached to the engine fan case, and is connected to the pylon by four bolts. Shear pins aid in aligning the mount with the pylon during engine installation. The forward engine mount supports side and vertical loads using a two-link arrangement with bolts loaded in shear. The main components of the forward mount assembly are the: •
Main beam
•
Side links
•
Shear pins
•
Failsafe bolt
The side and vertical loads at the front mount couple with the aft mount loads to support overall engine pitch and yaw. Side links provide the primary load paths from the fan case into the front beam. The front mount is attached to the pylon with four bolts and two shear pins, which transmit vertical and shear loads into the pylon. The mount bolts use captive barrel nuts to ease removal.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-42
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Fan Case
Fwd Mount Main Beam
Shear Pins
Fwd Mount Bolt Holes (4x)
Failsafe Bolt
PYLON Side Link Bolts (4x)
Fwd Mount Bolts (4x)
Captive Barrel Nuts FWD ENGINE MOUNT
CS1_CS3_7121_001
Side Link
Figure 21: Forward Engine Mount Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-43
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
AFT ENGINE MOUNT The aft engine mount is attached to the engine turbine exhaust case (TEC) and reacts to engine fore-aft, side, vertical, and roll loads. The main components of the aft mount assembly are: •
Main beam
•
Side links
•
Shear pins
•
Failsafe link
•
Balance beam
The two side links are assembled to the beam with four shear bolts. Each side link is attached to the turbine exhaust case with one shear bolt. The center link and the two attaching bolts are referred to as the failsafe link and bolts. THRUST LINKS Two thrust links attach the compressor intermediate case (CIC) to the aft mount through a balance beam. The thrust links transfer thrust loads to the aft mount. Any load differential in the two thrust links are balanced around the center pivot of the balance beam. The failure of any one link on the rear mount, including the thrust links, cause the system to transfer loads to a secondary load path.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-44
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Shear Pins Aft Mount Main Beam Thrust Links Failsafe Bolts
Balance Beam
Aft Mount Bolt Holes (4x)
Failsafe Link AFT ENGINE MOUNT
Side Links
CS1_CS3_7121_002
Aft Mount Bolts (4x)
Figure 22: Aft Engine Mount and Thrust Links Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-45
CS130_71_L3_EMP_DetailedComponentInformation.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
PRACTICAL ASPECTS ENGINE REMOVAL BOOTSTRAP EQUIPMENT The engine is removed using bootstrap equipment. The equipment consists of forward and aft bootstrap assemblies. Each assembly consists of arms and pylon mount attach points. Hoists, fitted to each arm, are used to raise and lower the engine cradle. The arms and pylon mount attach points are assembled and attached to the pylon fittings using quick-release pins.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-46
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
A
Forward Bootstrap Assembly Arms B
Quick-Release Pins
Hoist
A
B
CS1_CS3_7122_001
Aft Bootstrap Assembly Arms
Quick-Release Pins
Figure 23: Engine Removal Bootstrap Equipment Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-47
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
ENGINE CRADLE AND TRANSPORT STAND The engine cradle sits on a transport stand so that the engine can be moved. To remove the engine, the transport stand and cradle are positioned under the engine. The bootstrap hoists attach to a cradle that holds the engine. The cradle is raised up to the engine using the hoists and is attached to the engine. When the engine is disconnected from the pylon, the cradle and engine are lowered back onto the transport stand.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-48
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Hoist
Transport Stand
CS1_CS3_7122_002
Cradle
Figure 24: Engine Cradle, and Transport Stand Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-49
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
ELECTRICAL HARNESSES Electrical connectors are located at three positions on the forward pylon. There is a bracket with four connectors on each side of the engine fan case, and one bracket with three connectors on top of the engine near the precooler. All of these connectors need to be disconnected using soft jaw pliers prior to removing the engine. The generator power feeder harness connects to a terminal block, mounted near the precooler. The terminal block is accessed by removing the cover from the cooling conduit. The ring terminals need to be removed from the terminals and marked prior to removing the engine.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-50
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
D
A C
A RIGHT FAN ELECTRICAL DISCONNECTS
C VARIABLE FREQUENCY GENERATOR CABLE CONNECTS
B LEFT FAN ELECTRICAL DISCONNECTS
D VARIABLE FREQUENCY GENERATOR ELECTRICAL DISCONNECTS
CS1_CS3_7150_001
B
Figure 25: Electrical Harness Disconnects Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-51
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
HYDRAULIC AND FUEL CONNECTIONS A bracket mounted on the precooler inlet duct supports three hydraulic line quick-disconnect fittings, and two fuel line connections.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
71-52
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-40 Engine Mounts and Pylon
Fuel line Connections
CS1_CS3_7150_002
Hydraulic line Quick-Disconnect Fittings
Figure 26: Hydraulic and Fuel Connections Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-53
CS130_71_L2_EMP_PracticalAspects.fm
71 - Power Plant 71-70 Engine Drains
71-70 ENGINE DRAINS GENERAL DESCRIPTION The engine drain system collects and discharges oil, fuel, air, and hydraulic fluid from the engine through various tubes, forming a drain mast. The drain mast terminates at the latch access panel at the 6 o’clock position on the bottom of the engine cowling. The nacelle is also designed to drain fluid. Drain paths include gaps around the latch access panel, and drain holes in the lower bifurcation fixed/access panels. Components that drain through the drain mast include: •
2.5 bleed
•
High-pressure compressor (HPC) stator vane actuator
•
Low-pressure compressor (LPC) stator vane actuator
•
Variable frequency generator (VFG)
•
Starter
•
Fuel oil manifold and integrated fuel pump and control (IFPC)
•
Oil tank scupper
•
Hydraulic pump
A map of how the drain tubes are situated in the mast assists in troubleshooting. The map is located on the inside of the latch access panel door.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-54
CS130_71_L2_EngineDrains_General.fm
71 - Power Plant 71-70 Engine Drains
Fuel/Oil Manifold and IFPC
2.5 Bleed
HPC SVA
Hydraulic Pump
LPC SVA
A
B Starter C
Variable Frequency Generator
COWLING LATCH ACCESS PANEL
DRAIN TUBES MAP Oil Tank 1
DRAIN TUBES MAP DRAIN TUBES MAP 1. 999-5822-1 DR11, 2.5 BLEED 2. 999-5827-1 DR11, HPC-SVA
4
3 5
7
4. 999-5825-1 DR61, VFG 6
8
A
6. 999-5829-1 DR41, FOM-FUEL 7. 999-5826-1 DR71, OIL-TANK
3. 999-5823-1 DR01, LPC-SVA
LEGEND
5. 999-5832-1 DR91, STARTER
8. 999-5821-1 DR51, HYDRAULIC
B
DRAIN MAST
C
HPC IFPC LPC SVA
High-Pressure Compressor Integrated Fuel Pump and Control Low-Pressure Compressor Stator Vane Actuator
CS1_CS3_7171_001
2
Figure 27: Engine Drains Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-55
CS130_71_L2_EngineDrains_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
71-00 ENGINE STORAGE AND PRESERVATION GENERAL DESCRIPTION Turbine engines are designed to operate on a regular basis. Corrosion of vital engine parts can occur very quickly, especially in harsh environments such as coastal areas where the engine is exposed to salt air. Storage and preservation procedures are designed to minimize the way the environment may effect the engines. Engine covers must be installed on both engines to protect engine parts during shipping and handling, and from environmental conditions. Engine covers should always be installed if the engine is outdoors for extended periods without being operated.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-56
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
Inlet Cover
Right Rear Fan Case Cover
Hydraulic Pump Cover
Integrated Fuel Pump and Control Cover
Left Rear Fan Case Cover
Exhaust Cover
Air Turbine Starter Cover
Variable Frequency Generator Cover
CS1_CS3_7100_010
Deoiler Cover
Figure 28: Engine Storage and Preservation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-57
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
ENGINE PRESERVATION CHART
•
Check fuel for water using the ASTM D4176 procedure
The engines should be operated at least weekly. For engines that are not operated at least once a week, the following engine preservation guidelines are provided.
•
If required, drain fuel from the tanks until fuel passes the test procedure
Always refer to the aircraft maintenance publication (AMP) for preservation requirements and instructions. Method I (60 days or less) Whenever engine preservation is going to be 60 days or less the following steps need to be carried out: •
Clean and degrease the engine
•
Check for oil leaks
•
If components are removed from the engine or gearbox, ensure they are properly protected and preserved from corrosion
•
Sample the engine oil for water using ASTM D6304-00 procedure, if the sample fails then flush the engine oil system
•
Drain the engine oil system
•
Change the starter oil
•
Install tarp or PVC plastic sheets over inlet cowl, fan exhaust area and engine exhaust area
•
Make a record of the engine preservation
There is a provision to extend the engine preservation an extra 60 days if the following procedures are carried out: •
Remove the engine covers
•
Service the engine oil system
•
Run the engine at idle for five minutes once the oil temperature is above 220 °F (104.4 °C)
•
Shut down the engine
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Method II (More than 60 days) Whenever engine preservation is going to extend beyond 60 days, the following steps need to be carried out: •
Drain the engine oil system
•
Replace engine oil filter and fill the oil system
•
Start the engine and run at idle or dry motor for 3 minutes using the engine starter. It is necessary that the fan turn so that oil reaches through all of the engine oil system
•
Preserve the fuel system using preservation oil
•
Drain the engine oil
•
Change the starter oil
•
Clean and preserve the gearbox
•
Place some desiccant and some relative humidity indicators inside engine covers
•
Install tarp or PVC plastic sheets over inlet cowl, fan exhaust area and engine exhaust area. Make sure the relative humidity indicators are visible through windows in the protective covers
•
Make a record of the engine preservation
TECHNICAL TRAINING MANUAL
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71-58
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
Determine engine preservation time
More than 60 days
60 days or less
1 - 7 days
or
No action necessary
Option A
Option B
Operate the engine
Preserve for 60 days or less by (METHOD I)
Day 8
Sample the engine oil by ASTM D6304-00 procedure Preserve for 60 days or less by (METHOD I) Do not flush and drain the oil
Results OK
Results not OK
Resample the engine oil by ASTM D6304-00 procedure
Day 61
Results not OK
Do option A Run engine (option available when (METHOD I) is repeated a maximum of 1 time)
Day 121
Preserve for more than 60 days by (METHOD II) NOTE 1. This chart does not specify all of the steps in the preservation method. 2. Sampling option is restricted to one 60 day period. 3. In option B, oil system flush is not necessary (ref. METHOD I) if sample test results are OK. 4. When preserving for more than 60 days, fuel system preservation is also necessary. 5. If the engine remains in storage after 60 days (METHOD I), preserve the engine again after the initial 60 day period.
CS1_CS3_7100_011
Return engine to service
Figure 29: Engine Preservation Chart Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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71-59
CS130_71_L2_ESP_General.fm
71 - Power Plant 71-00 Engine Storage and Preservation
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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71-60
CS130_71_ILB.fm
ATA 72 - Engine
BD-500-1A10 BD-500-1A11
72 - Engine Fuel and Control
Table of Contents 72-00 PW1500G Engine.........................................................72-2 General Description ..........................................................72-2 Component Location ........................................................72-4 72-11 Fan Assembly...............................................................72-6 General Description ..........................................................72-6 Detailed Component Information ......................................72-8 Fan Hub .......................................................................72-8 Fan Blades...................................................................72-8 Inlet Cone.....................................................................72-8 Fan Case Assembly...................................................72-10 72-15 Fan Drive Gear System ..............................................72-12 General Description ........................................................72-12 Detailed Component Information ....................................72-14 Fan Drive Gearbox.....................................................72-14 No. 1 and No. 1.5 Bearing Support Assembly ...........72-16 72-22 Fan Intermediate Case ...............................................72-18 General Description ........................................................72-18 Detailed Component Information ....................................72-20 Fan Exit Fairing..........................................................72-20 Fan Exit Stator ...........................................................72-20 Fan Intermediate Case ..............................................72-20 No.2 Bearing and Support Assembly.........................72-20 Fan Drive Gearbox Coupling .....................................72-20 Low-Pressure Compressor Variable Stator Vanes ...72-20 72-31 Low-Pressure Compressor.........................................72-22 General Description ........................................................72-22 Detailed Component Information ....................................72-24 Low-Pressure Compressor Case...............................72-24 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Low-Pressure Compressor Stators ........................... 72-24 Integral Bladed Rotor ................................................ 72-24 No.2 Bearing ............................................................. 72-24 Tie Rods.................................................................... 72-24 72-34 Compressor Intermediate Case ................................. 72-26 General Description ....................................................... 72-26 Detailed Component Information.................................... 72-28 Compressor Intermediate Case ................................ 72-28 No. 3 Bearing Compartment ..................................... 72-28 Gearbox Drive Bevel Gear ........................................ 72-28 Fire Shield ................................................................. 72-28 Low-Pressure Compressor Exit Stator ...................... 72-28 High-Pressure Compressor Variable Inlet Guide Vanes ....................................... 72-28 72-35 High-Pressure Compressor ....................................... 72-30 General Description ....................................................... 72-30 Detailed Component Information ................................... 72-32 Rotors........................................................................ 72-32 High-Pressure Compressor Rotor Shaft ................... 72-32 Stators....................................................................... 72-32 72-41 Diffuser, Combustor and Turbine Nozzle Assembly .................................................... 72-34 General Description ....................................................... 72-34 Detailed Component Information ................................... 72-36 Diffuser Case Assembly............................................ 72-36 Combustion Chamber ............................................... 72-36 High-Pressure Turbine Nozzle Assembly ................. 72-36 72-51 High-Pressure Turbine............................................... 72-38 General Description ....................................................... 72-38
TECHNICAL TRAINING MANUAL
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72-i
CS130_M5_72_EngineTOC.fm
72 - Engine Fuel and Control
Detailed Component Information ....................................72-40 High-Pressure Turbine...............................................72-40 72-52 Turbine Intermediate Case .........................................72-42 General Description ........................................................72-42 Detailed Component Information ....................................72-44 Turbine Intermediate Case ........................................72-44 72-53 Low-Pressure Turbine ................................................72-46 General Description ........................................................72-46 Detailed Component Information ....................................72-48 Low-Pressure Turbine................................................72-48 72-54 Turbine Exhaust Case ................................................72-50 General Description ........................................................72-50 Detailed Component Information ....................................72-52 Turbine Exhaust Case ...............................................72-52 72-60 Main and Angle Gearboxes ........................................72-54 General Description ........................................................72-54 Main Gearbox ............................................................72-54 Main Gearbox Mounted Components ........................72-56 Angle Gearbox ...........................................................72-58 72-00 Borescope Access......................................................72-60 General Description ........................................................72-60 Engine Left Side Borescope Access..........................72-60 Engine Right Side Borescope Access .......................72-62 Borescope Access - Main Gearbox ...........................72-64 Borescope Access - Angle Gearbox ..........................72-64
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TECHNICAL TRAINING MANUAL
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72-ii
CS130_M5_72_EngineTOC.fm
72 - Engine Fuel and Control
List of Figures Figure 1: PW1500G Engine..................................................72-3 Figure 2: PW1500G Engine Assembly Component Location .............................................72-5 Figure 3: Fan Assembly........................................................72-7 Figure 4: Fan Hub, Blades, and Inlet Cone ..........................72-9 Figure 5: Fan Case Assembly ............................................72-11 Figure 6: Fan Drive Gear System.......................................72-13 Figure 7: Fan Drive Gearbox ..............................................72-15 Figure 8: No. 1 and No. 1.5 Bearing Support Assembly and Cross-Section ..............................72-17 Figure 9: Fan Intermediate Case........................................72-19 Figure 10: Fan Intermediate Case Cross-Section ...............72-21 Figure 11: Low-Pressure Compressor..................................72-23 Figure 12: Low-Pressure Compressor Cross-Section .........72-25 Figure 13: Compressor Intermediate Case ..........................72-27 Figure 14: Compressor Intermediate Case Cross-Section......................................................72-29 Figure 15: High-Pressure Compressor.................................72-31 Figure 16: High-Pressure Compressor Cross-Section .........72-33 Figure 17: Diffuser, Combustor, and Turbine Nozzle Assembly....................................72-35 Figure 18: Diffuser and Combustor Cross-Section...............72-37 Figure 19: High-Pressure Turbine .......................................72-39 Figure 20: High-Pressure Turbine Cross-Section.................72-41 Figure 21: Turbine Intermediate Case.................................72-43 Figure 22: Turbine Intermediate Case Cross-Section ..........72-45 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 23: Low-Pressure Turbine ........................................ 72-47 Figure 24: Low-Pressure Turbine Cross-Section................. 72-49 Figure 25: Turbine Exhaust Case ....................................... 72-51 Figure 26: Turbine Exhaust Case Cross-Section................. 72-53 Figure 27: Main Gearbox ..................................................... 72-55 Figure 28: Main Gearbox Mounted Components................. 72-57 Figure 29: Angle Gearbox.................................................... 72-59 Figure 30: Engine Left Side Borescope Access................... 72-61 Figure 31: Engine Right Side Borescope Access ................ 72-63 Figure 32: Borescope Access Gearbox and Angle Gearbox .............................. 72-65
TECHNICAL TRAINING MANUAL
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72-iii
CS130_M5_72_EngineLOF.fm
ENGINE - CHAPTER BREAKDOWN ENGINE CHAPTER BREAKDOWN
Engine
1
72 - Engine 72-00 PW1500G Engine
72-00 PW1500G ENGINE GENERAL DESCRIPTION The PW1500G series engines are high bypass ratio turbofans, utilizing two spools with an integral primary and fan duct exhaust systems. The N1 rotor consists of a single-stage low-pressure ratio fan, driven at a reduced speed through a fan drive gear system (FDGS), a 3-stage low-pressure compressor (LPC), and a 3-stage low-pressure turbine (LPT) on a common shaft. The FDGS is a planetary star gear reduction unit that takes torque from the LPT and reduces the fan speed to 30% of the LPT speed. The LPC has variable inlet guide vanes, which are positioned automatically by the electronic engine control (EEC) to optimize engine operation. The N2 rotor consists of the high-pressure compressor (HPC) with eight compressor stages, and a high-pressure turbine (HPT) consisting of two stages. The HPC has variable inlet guide vanes and variable stator vanes in the 1st, 2nd, and 3rd stages which are positioned automatically by the EEC to optimize engine operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-2
CS130_72_L2_PW1500G_General.fm
72 - Engine 72-00 PW1500G Engine
3 Stage Low-Pressure Compressor
8 Stage High-Pressure Compressor
2 Stage High-Pressure Turbine
3 Stage Low-Pressure Turbine
Fan Rotor
2
2.5
3
4
4.5 5
1
LEGEND NOTE
Fan Rotor/FDGS N1 Rotor N2 Rotor
1
Engine station numbers.
CS1_CS3_7200_009
4.9
Figure 1: PW1500G Engine Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-3
CS130_72_L2_PW1500G_General.fm
72 - Engine 72-00 PW1500G Engine
COMPONENT LOCATION The PW1500G engine is made up of the following modules: •
Fan rotor
•
Fan case
•
Fan drive gear system (FDGS)
•
Fan intermediate case (FIC)
•
Low-pressure compressor (LPC)
•
Compressor intermediate case (CIC)
•
High-pressure compressor (HPC)
•
Diffuser and combustor
•
High-pressure turbine (HPT)
•
Turbine intermediate case (TIC)
•
Low-pressure turbine (LPT)
•
Turbine exhaust case (TEC)
•
Angle gearbox
•
Main gearbox (MGB)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-4
CS130_72_L2_PW1500G_ComponentLocation.fm
72 - Engine 72-00 PW1500G Engine
Fan Case
FDGS
CIC DIFF/COMB FDGS FIC HPC HPT LPC LPT TEC TIC
Compressor Intermediate Case Diffuser/Combustor Fan Drive Gear System Fan Intermediate Case High-Pressure Compressor High-Pressure Turbine Low-Pressure Compressor Low-Pressure Turbine Turbine Exhaust Case Turbine Intermediate Case
Angle Gearbox
FAN ROTOR
FIC
LPC
CIC
Main Gearbox
HPC
DIFF/ COMB
ENGINE MODULES
HPT
TIC
LPT
TEC
CS1_CS3_7200_010
LEGEND
Figure 2: PW1500G Engine Assembly Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-5
CS130_72_L2_PW1500G_ComponentLocation.fm
72 - Engine 72-11 Fan Assembly
72-11 FAN ASSEMBLY GENERAL DESCRIPTION The fan assembly is composed of the fan rotor and the fan case assembly. The fan rotor is located inside the fan case assembly. It turns clockwise and produces 90% of the engine total thrust. The fan case assembly directs the fan air through and around the engine core.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-6
CS130_72_L2_FanAssembly_General.fm
72 - Engine 72-11 Fan Assembly
Fan Case
CS1_CS3_7200_008
Fan Rotor
Figure 3: Fan Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-7
CS130_72_L2_FanAssembly_General.fm
72 - Engine 72-11 Fan Assembly
DETAILED COMPONENT INFORMATION FAN HUB The fan hub retains the fan blades on its circumference. The hub has attachment points on the front face for the inlet cone and fan rotor trim balance weights. The no. 1 fan blade slot is located between two no.1 marks on the hub. Numbering of the fan blades and fairings is in the direction of rotation, beginning at the no. 1 blade position. Details on fan rotor balancing can be found in ATA 77, Engine Indicating. FAN BLADES The fan rotor has 18 wide-chord hollow aluminum fan blades. The hollow aluminum fan blades have titanium leading edges for wear resistance and foreign object debris (FOD) tolerance. A protective polyurethane black paint coating provides erosion protection for the fan blade airfoil. Each fan blade has a dovetail shaped root that fits into slots on the fan hub. Composite wear strips are bonded to the fan blade dovetail sides. Composite fan blade fairings are located between the fan blades to provide an inner flow path surface. Each fan blade fairing is held in place by a fan blade retaining pin, inserted into pin lugs on the fan hub. A spacer is installed between each blade and the hub to minimize wear, and provide a tight fit. The fan blades are held in place on the hub by a front locking ring. INLET CONE The composite inlet cone is installed in front of the fan hub to smooth the airflow into the engine. The inlet cone cover provides access to the inlet cone retaining hardware. The inlet cone is removed to gain access to the fan blades lock ring.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-8
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-11 Fan Assembly
Titanium Leading Edge
Aluminum Fan Blade
Fan Hub Fan Blade Fairing
Composite Wear Strip
Fan Blade Fairing Pin
Fan Blade Spacer
Lock Ring
1
Fan Blade Fairing Pin Lugs
Inlet Cone 1
Fan Rotor Trim Balance Weights Attachment Flange
Inlet Cone Attachment Flange
CS1_CS3_7211_002
No. 1 Fan Blade Marks
Inlet Cone Cover
Figure 4: Fan Hub, Blades, and Inlet Cone Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-9
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-11 Fan Assembly
FAN CASE ASSEMBLY
Precooler Duct Inlet
The fan case serves as the structural link between the inlet cowl and the engine core by providing an attachment point to the airframe. It consists of a one-piece composite case with a titanium mount ring attached around the aft outer surface. A Kevlar containment ring surrounds the fan case.
The precooler duct inlet is a four-piece titanium casting mounted at the rear top of the fan case. It is used to direct fan air into the pylon mounted precooler, provides seal lands for the thrust reverser (T/R) door firewall seal, and sealing of the upper nacelle flow path. The precooler duct inlet is attached to the titanium mount ring at the rear of the case. The mount ring provides the front mount attachment points for the engine mounts, and includes a V-groove around the entire circumference, providing alignment and support for the thrust reverser doors.
Fan case components include: •
Fan blade rub strips
•
Acoustic panels
•
Ice liners
•
Precooler duct inlet
•
Fan exit guide vanes (FEGV)
•
Fan exit liner segments
Fan Exit Guide Vanes On the inner surface of the fan case, 44 hollow aluminum fan exit guide vanes (FEGVs) extend rearward to the fan intermediate case (FIC). The FEGVs straighten the fan air and provide structural support between the FIC and the fan case assembly. A polyurethane coating covers most of the FEGV surfaces to protect against corrosion and erosion.
Fan Blade Rub Strips
Fan Exit Liner Segments
The fan blade rub strips provide an abradable surface for the fan blade to rub against when the blades elongate under power. They also protect the fan blades from contacting the fan case and ensure a tight clearance between the case and the rub strips. The surface of the rub strips can be refurbished if any significant rubbing occurs.
Six fan exit liner segments, immediately aft of the FEGVs, surround the low-pressure compressor (LPC). Stationary louvers in the fan exit liner segments provide an exit path for the 2.5 bleed air into the fan air stream.
Acoustic Panels
The segments can be removed individually for component access and inspection of lines and fittings.
Acoustic panels, forward of the fan rub strip, and acoustic liner segments aft of the fan blades, decrease noise generated from the fan. Ice Liners Replaceable ice liners aft of the fan, provide protection from ice for the case structure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-10
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-11 Fan Assembly
Fan Case
Engine Mount
Precooler Duct Inlet A
Fan Exit Liner Segments
A FORWARD ENGINE MOUNT V-Groove
Fan Blade Rub Strip Acoustic Panels
Ice Liner Rear Acoustic Liner Segment
Titanium Mount Ring
B FAN EXIT GUIDE VANE
CS1_CS3_7221_001
B
Figure 5: Fan Case Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-11
CS130_72_L3_FanAssembly_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
72-15 FAN DRIVE GEAR SYSTEM GENERAL DESCRIPTION The fan drive gear system (FDGS) is a reduction gear assembly that uses torque from the low-pressure turbine (LPT) N1 to turn the fan assembly at approximately one third the speed of the LPT. This results in a lower fan speed and a higher low-pressure compressor (LPC) speed, thus increasing compressor efficiency. The FDGS is located between the fan rotor assembly and the fan intermediate case (FIC). It is radially and axially supported by the no.1 and no. 1.5 bearing support assembly. The LPC shaft connects to the input of the FDGS. The input speed is reduced by the fan drive gearbox (FDG) and then transmitted to a gearbox shaft connected to the fan rotor hub.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-12
CS130_72_L2_FDGS_General.fm
72 - Engine 72-15 Fan Drive Gear System
Support Assemblies Fan Drive Gear System B
Input from Low-Pressure Turbine Shaft A
Fan Drive Gearbox
A
No. 1 and No. 1.5 Bearing Support
B FAN DRIVE GEAR SYSTEM
CS1_CS3_7210_001
Gearbox Shaft
Figure 6: Fan Drive Gear System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-13
CS130_72_L2_FDGS_General.fm
72 - Engine 72-15 Fan Drive Gear System
DETAILED COMPONENT INFORMATION FAN DRIVE GEARBOX The fan drive gearbox (FDG) is the most aft section of the FDGS. It houses the sun gear, the five star gears, and the rotating ring gear halves. The FDG takes the torque from the low-pressure turbine (LPT) FDG titanium coupling to turn a sun gear. This sun gear turns five stationary star gears against an outer ring gear set. The outer ring gear set is turned by the star gears and is connected to the gearbox shaft, which connects to the fan hub. Each star rotates around its own journal bearing mounted in the carrier. The journal bearings provide the rotational hinge for the star gears. The oil manifold supplies oil to no. 1 and no. 1.5 bearing support assembly, which supplies the lubrication for the FDG gears and bearings. The torque frame links the carrier to the flex support that is bolted to the fan intermediate case (FIC). This flex support ensures the FDGS remains aligned with the FDG coupling. The FDG coupling connects the LPC shaft to the FDGS. The FDG coupling is a specially designed flex shaft that is used to reduce pressure on the FDGS. It transmits the torque from the LPC to the FDGS input.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-14
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
Ring Gear Set Halves
Gearbox Shaft
Journal Bearings
Sun Gear
Torque Frame
FDG Coupling Gearbox Shaft No. 1 and No. 1.5 Bearing Support Assembly
Star Gears
Carrier
Oil Manifold
Flex Support
CS1_CS3_7210_003
Fan Drive Gearbox
Figure 7: Fan Drive Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-15
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
NO. 1 AND NO. 1.5 BEARING SUPPORT ASSEMBLY The no. 1 and no. 1.5 bearing support assembly supports the gearbox shaft using two tapered roller bearings. The no. 1 bearing is oil damped to absorb fan vibration. The aft section of the bearing support assembly is attached to the fan intermediate case (FIC).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-16
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-15 Fan Drive Gear System
No. 1 Bearing Support No. 1 and No. 1.5 Bearing Support
Gearbox Shaft
No. 1.5 Bearing CS1_CS3_7215_001
No. 1 Bearing
Figure 8: No. 1 and No. 1.5 Bearing Support Assembly and Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-17
CS130_72_L3_FDGS_DetailedComponentInfo.fm
72 - Engine 72-22 Fan Intermediate Case
72-22 FAN INTERMEDIATE CASE GENERAL DESCRIPTION The fan intermediate case (FIC) provides the support structure between the fan case, the fan drive gear system (FDGS), and the low-pressure compressor (LPC). The fan drive gearbox (FDG) is housed inside the FIC. It also provides the aerodynamic flow path between the fan rotor and the LPC. The fan exit fairing splits the air between the fan duct and the low-pressure compressor. Air entering the fan exit stators is directed to the LPC variable stator vanes.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-18
CS130_72_L2_FIC_General.fm
72 - Engine 72-22 Fan Intermediate Case
Fan Intermediate Case
Fan Exit Stator
CS1_CS3_7222_011
Fan Exit Fairing
Figure 9: Fan Intermediate Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-19
CS130_72_L2_FIC_General.fm
72 - Engine 72-22 Fan Intermediate Case
DETAILED COMPONENT INFORMATION
LOW-PRESSURE COMPRESSOR VARIABLE STATOR VANES
FAN EXIT FAIRING
The low-pressure compressor (LPC) variable stator vanes are located at the FIC aft end and mate with the LPC case front end. The LPC variable stators vane improve engine stability and operation.
The fan exit fairing is installed on the front flange of the fan intermediate case (FIC). It splits the airflow between the engine core section, and the fan. FAN EXIT STATOR The fan exit stator is installed on the front flange of the FIC, just inside of the splitter. It straightens the fan air flow and directs it to the low-pressure compressor (LPC) variable vane stator. FAN INTERMEDIATE CASE The fan intermediate case (FIC) keeps the fan drive gear system (FDGS) and the LPC in alignment, and also provides support for the no. 1 and no. 1.5 bearing support assembly. NO.2 BEARING AND SUPPORT ASSEMBLY The no. 2 bearing and support assembly is attached to the rear center of the FIC. It absorbs rotor radial movement, and is oil damped on the outer race to decrease vibration. FAN DRIVE GEARBOX COUPLING The fan drive gearbox coupling (FDG coupling) is located in the heart of the FIC. Its aft end sits inside of bearing no. 2 and extends forward to couple with the fan drive gearbox system (FDGS).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-20
CS130_72_L3_FIC_DetailedComponentInfo.fm
72 - Engine 72-22 Fan Intermediate Case
Fan Exit Fairing
Fan Intermediate Case
Fan Exit Stator
Low-Pressure Compressor Variable Stator Vane
Low-Pressure Compressor
No. 1 and No. 1.5 Bearing Support
No. 2 Bearing and Assembly Support
Low-Pressure Compressor Rotor Hub
CS1_CS3_7222_001
Fan Drive Gearbox Coupling
Figure 10: Fan Intermediate Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-21
CS130_72_L3_FIC_DetailedComponentInfo.fm
72 - Engine 72-31 Low-Pressure Compressor
72-31 LOW-PRESSURE COMPRESSOR GENERAL DESCRIPTION The low-pressure compressor (LPC) has three stages. Each stage has an integral bladed rotor (IBR). An IBR consists of a disc and blades manufactured as a single unit, resulting in increased efficiency and less weight than conventional bladed discs. An annular bleed valve at the rear of the LPC compressor case, designated as the 2.5 bleed valve, provides operational stability for the LPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-22
CS130_72_L2_LPC_General.fm
72 - Engine 72-31 Low-Pressure Compressor
Low-Pressure Compressor Case
2.5 Bleed Valve
CS1_CS3_7231_001
Stage 1 Integral Bladed Rotor (IBR)
Figure 11: Low-Pressure Compressor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-23
CS130_72_L2_LPC_General.fm
72 - Engine 72-31 Low-Pressure Compressor
DETAILED COMPONENT INFORMATION LOW-PRESSURE COMPRESSOR CASE The 2.5 bleed valve is mounted on the LPC case. Removable fan exit liner segments cover the LPC case. LOW-PRESSURE COMPRESSOR STATORS There are two stages of stator vanes in the LPC. Each stage consists of fixed aluminum segment stators. INTEGRAL BLADED ROTOR The three integral bladed rotors (IBRs) are tied to the LPC rotor hub outer diameter using tie rods. The LPC rotor hub is splined onto the low-pressure turbine (LPT) shaft, just aft of the oil damped no. 2 bearing. It also couples to the fan drive gearbox (FDG) coupling. NO.2 BEARING The no. 2 bearing supports the LPC rotor hub, the FDG coupling, and the LPT shaft. TIE RODS There are 48 tie rods that connect the IBRs to the LPC rotor hub. They are located on the second IBR.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-24
CS130_72_L3_LPC_DetailedComponentInfo.fm
72 - Engine 72-31 Low-Pressure Compressor
Fan Exit Liner Segment
Low-Pressure Compressor Case
Stator Vanes
2.5 Bleed Valve
Integral Bladed Rotors
Tie Rod
No. 2 Bearing
Low-Pressure Compressor Rotor Hub
Low-Pressure Turbine Shaft
CS1_CS3_7231_002
Fan Drive Gearbox Coupling
Figure 12: Low-Pressure Compressor Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-25
CS130_72_L3_LPC_DetailedComponentInfo.fm
72 - Engine 72-34 Compressor Intermediate Case
72-34 COMPRESSOR INTERMEDIATE CASE GENERAL DESCRIPTION The compressor intermediate case (CIC) acts as the support structure for the low-pressure compressor (LPC), and the high-pressure compressor (HPC). The CIC carries the thrust loads from the engine through the two thrust links attachment points to the aft engine mounts. The two thrust link attachment points are on the CIC at the 2 o’clock and 10 o’clock positions. The CIC also houses the no. 3 bearing compartment. The case structure includes nine struts and an opening for a gearbox drive shaft that transmits torque to the angle gearbox. The case is equipped with an opening for the discharge of the 2.5 bleed air. The hollow struts provide support for the no.3 bearing compartment. The struts interior provide access to the no.3 bearing for lubrication supply and scavenge lines. The struts are numbered clockwise beginning at the 1 o’clock position. A table lists all the struts internal passage usage.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-26
CS130_72_L2_CIC_General.fm
72 - Engine 72-34 Compressor Intermediate Case
Strut 9
Strut 1
Strut 8
High-Pressure Compressor Variable Inlet Guide Vanes
Strut 2
Strut 7
Strut 3
Strut 6
Strut 4
Thrust Link Attachment Point
Strut 5 VIEW LOOKING FORWARD
COMPRESSOR INTERMEDIATE CASE STRUTS
Low-Pressure Compressor Exit Stator
No. 3 Bearing Compatment
2.5 Bleed Air Openings
Used For
1
No. 2 bearing buffer air supply
2
Breather tube
3
No. 3 bearing oil supply
4
N1 probe
5
No. 3 bearing oil scavenge
6
Gearbox drive shaft
7
Empty
8
No. 3 bearing buffer air supply
9
No. 3 bearing buffer air supply
CS1_CS3_7234_001
Strut No.
Figure 13: Compressor Intermediate Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-27
CS130_72_L2_CIC_General.fm
72 - Engine 72-34 Compressor Intermediate Case
DETAILED COMPONENT INFORMATION
HIGH-PRESSURE COMPRESSOR VARIABLE INLET GUIDE VANES
COMPRESSOR INTERMEDIATE CASE
The high-pressure compressor variable inlet guide vanes are located at the aft end of the CIC. They are adjusted during engine operation to optimize engine performance.
The compressor intermediate case (CIC) includes the no. 3 bearing, the 2.5 bleed air outlet, and the thrust reverser (T/R) door inner fixed structure (IFS) support. NO. 3 BEARING COMPARTMENT The no. 3 bearing compartment is located in the aft section of the compressor intermediate case (CIC). The no. 3 bearing supports the front of the high-pressure compressor’s (HPC) rotor shaft axial movement and absorbs vibrations. GEARBOX DRIVE BEVEL GEAR The gearbox drive bevel gear is located at the 8 o’clock position in the no. 3 bearing compartment. The no. 3 bearing bevel gear engages the gearbox drive bevel gear to supply torque to the angle gearbox drive shaft. This shaft drives the angle gearbox, which in turn, supplies torque to the main gearbox. FIRE SHIELD The fire shield is attached to the outer perimeter of the CICs aft section. It is a metal containment shield used to protect the LPC structure from fire originating in the engine core compartment. The outer section of the shield acts as a support for the thrust reverser (T/R) door inner fixed structure (IFS). LOW-PRESSURE COMPRESSOR EXIT STATOR The low-pressure compressor (LPC) exit stator segments are located at the front of the CIC. The LPC exit stators divert air from the LPC, and direct it towards the HPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-28
CS130_72_L3_CIC_DetailedComponentInfo.fm
72 - Engine 72-34 Compressor Intermediate Case
Gearbox Drive Bevel Gear
No. 3 Bearing Bevel Gear
No. 3 Bearing Compartment
High-Pressure Compressor Rotor Shaft Angle Gearbox Drive Shaft Gearbox Drive Bevel Gear Support
High-Pressure Compressor High-Pressure Compressor Variable Inlet Guide Vanes
Low-Pressure Compressor No. 2.5 Exit Stators Bleed Air Outlet
Fire Shield
Angle Gearbox Thrust Reverser Door Inner Fixed Structure Support
CS1_CS3_7234_003
Low-Pressure Compressor
Figure 14: Compressor Intermediate Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-29
CS130_72_L3_CIC_DetailedComponentInfo.fm
72 - Engine 72-35 High-Pressure Compressor
72-35 HIGH-PRESSURE COMPRESSOR GENERAL DESCRIPTION The high-pressure compressor (HPC) increases the speed and pressure of the primary gaspath air. The HPC is driven by the high-pressure turbine (HPT) through the HPC rotor shaft. The first seven rotor stages are integral bladed rotors (IBRs). The first three stator stages are variable and actuated through unison rings on the outside of the HPC case. The forward HPC case is split horizontally and mated with a one piece aft ring case. The HPC supplies 4th and 8th stage air to the bleed air system. The 4th stage bleed air port is located on the HPC. Engine cooling is supplied by 4th stage air through cooling air ports on the HPC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-30
CS130_72_L2_HPC_General.fm
72 - Engine 72-35 High-Pressure Compressor
4th Stage Bleed Air Port
High-Pressure Compressor Rotor Shaft
Stator Vane Unison Rings
Aft Ring Case
Split Front Case Stage 1 Integral Bladed Rotor
CS1 CS3 7235 001 CS1_CS3_7235_001
4th Stage Cooling Air Ports
Figure 15: High-Pressure Compressor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-31
CS130_72_L2_HPC_General.fm
72 - Engine 72-35 High-Pressure Compressor
DETAILED COMPONENT INFORMATION ROTORS There are eight rotor stages in the high-pressure compressor (HPC). The first seven rotor stages are integral bladed rotors (IBRs). The 8th stage rotor is a standard blade and disk assembly. HIGH-PRESSURE COMPRESSOR ROTOR SHAFT The rotor stage hubs are held together on the HPC rotor shaft. The HPC front hub acts as the forward stop for the IBRs and the HPC rear hub secures the 8th stage rotor. This rotor hub is held by the HPC rotor shaft nut on the HPC rotor shaft. STATORS The first three stator stages of the HPC are variable, while the remaining four stages are fixed stages.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-32
CS130_72_L3_HPC_DetailedComponentInfo.fm
72 - Engine 72-35 High-Pressure Compressor
HPC Front Hub
Integral Bladed Rotors
Variable Stator Vanes
8th Stage Rotor
Fixed Stator Vanes
HPC Rear Hub
LEGEND HPC
HPC Rotor Shaft
HPC Rotor Shaft Nut
CS1_CS3_7235_002
Compressor Intermediate Case
High-Pressure Compressor
Figure 16: High-Pressure Compressor Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-33
CS130_72_L3_HPC_DetailedComponentInfo.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
72-41 DIFFUSER, COMBUSTOR AND TURBINE NOZZLE ASSEMBLY GENERAL DESCRIPTION The diffuser case, combustion chamber, and turbine nozzle assemblies work together to provide the internal combustion that drives the turbines. Compressed air flows from the high-pressure compressor (HPC) exit stator into the diffuser and then into the combustor section. The airflow is straightened and diffused, reducing the velocity, while maintaining pressure before it enters the combustion chamber. The diffuser case has the bleed air ports for the 6th stage bleed air, cooling air, and cowl anti-ice (CAI). The combustion chamber inner perimeter and the turbine nozzles are attached to the diffuser inner case. The combustion chamber consists of an inner and outer chamber liner assembly. The diffuser case also has provisions for fuel nozzles and igniter plugs. Metered fuel is supplied to the combustion chamber through 16 fuel nozzles, where it is mixed with the air. The heat of the combustion process causes the airflow to expand and accelerate rearward. Turbine nozzle guide vanes then direct the high-temperature, high-velocity gases out of the combustion chamber to drive the turbines. The diffuser outer case supports the main gearbox (MGB) using two mount brackets, located at the 3 o'clock and 9 o'clock position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-34
CS130_72_L2_DCTNA_General.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
Turbine Nozzle Assembly
Inner Liner
Outer Liner COMBUSTION CHAMBER
Diffuser Outer Case High-Pressure Compressor Exit Stator Gearbox Mount Bracket (2x)
Diffuser Inner Case DIFFUSER CASE ASSEMBLY
6th Stage Cooling Air Port
CS1_CS3_7240_001
Fuel Nozzle (16x)
Figure 17: Diffuser, Combustor, and Turbine Nozzle Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-35
CS130_72_L2_DCTNA_General.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
DETAILED COMPONENT INFORMATION DIFFUSER CASE ASSEMBLY The diffuser inner case attaches to a common diffuser outer case flange with the high-pressure compressor (HPC). The inner case attaches to the high-pressure turbine (HPT) flange. Two igniter plugs located at the 4 o’clock and 5 o'clock position extend into the combustion chamber. COMBUSTION CHAMBER The HPC air enters the diffuser and flows through the hood and cooling holes around the inner and outer liners of the combustion chamber. A bulkhead is used to attach both combustion chamber inner and outer liners, and the hood. Fuel nozzles inject fuel through orifices in the bulkhead where swirlers spin the air/fuel mixture inside the combustion chamber. HIGH-PRESSURE TURBINE NOZZLE ASSEMBLY The turbine nozzle assembly attaches to the cooling air duct, which is bolted with the combustion chamber inner liner flange, to the inner diffuser case. The nozzle assembly has 32 air-cooled thermal-barrier coated guide vanes between the inner and outer combustion chamber liners. Located at the rear of the combustion chamber, they direct the hot gases to the 1st stage high-pressure turbine blades. Each hollow vane is cooled by diffuser air and use cooling air exit holes on the airfoil external surface to reduce vane temperature.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-36
CS130_72_L3_DCTNA_DetailedComponentInfo.fm
72 - Engine 72-41 Diffuser, Combustor and Turbine Nozzle Assembly
Fuel Nozzle
High-Pressure Compressor
Igniter Plug
Diffuser Case High-Pressure Compressor Exit Stator
High-Pressure Turbine
Outer Liner
Hood Combustion Chamber
1st Stage Vane Bulkhead Swirler
Inner Liner
CS1_CS3_7241_001
Inner Diffuser Case
Cooling Air Holes
Figure 18: Diffuser and Combustor Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-37
CS130_72_L3_DCTNA_DetailedComponentInfo.fm
72 - Engine 72-51 High-Pressure Turbine
72-51 HIGH-PRESSURE TURBINE GENERAL DESCRIPTION The high-pressure turbine (HPT) provides the rotational force to drive the HPC by extracting energy from the hot combustion gases. The HPT case provides containment of the HPT rotors. Cooling air flows through the hollow 2nd stage stator vanes to keep them cool.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-38
CS130_72_L2_HPT_General.fm
72 - Engine 72-51 High-Pressure Turbine
CS1_CS3_7251_001
2nd Stage Vane Cooling Air
Figure 19: High-Pressure Turbine Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-39
CS130_72_L2_HPT_General.fm
72 - Engine 72-51 High-Pressure Turbine
DETAILED COMPONENT INFORMATION HIGH-PRESSURE TURBINE Each of the two turbine rotor stages contain 44 blades coated with a thermal barrier and internally cooled. The 1st stage hub engages a spline in the HPC rear hub at the front. The 2nd stage hub is attached by a nut on the HPC rotor shaft at the rear. The blades are attached to their hubs in fir tree type slots and held in position with retaining plates. The blades outer air seal segments are ground as a set with the turbine blades to maximize rotor tip sealing. Internally cooled 2nd stage stator vanes are mounted inside the case between the two rotor stages. Fan air from the active clearance control (ACC) system cools the HPT case, thus decreasing its inner diameter and increasing the turbine efficiency. The HPT case has mounting points for the turbine ACC manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-40
CS130_72_L3_HPT_DetailedComponentInfo.fm
72 - Engine 72-51 High-Pressure Turbine
HPT Case
Turbine Active Clearance Control (ACC) Manifold Mount
HPT 2nd Stage Vane
Turbine Rotor Stages
HPC Rotor Shaft
HPT 1st Stage Hub
HPT 2nd Stage Hub
CS1_CS3_7251_002
HPC Rear Hub
Figure 20: High-Pressure Turbine Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-41
CS130_72_L3_HPT_DetailedComponentInfo.fm
72 - Engine 72-52 Turbine Intermediate Case
72-52 TURBINE INTERMEDIATE CASE GENERAL DESCRIPTION The turbine intermediate case (TIC) is used to turn the high-pressure turbine (HPT) airflow to align it with the counter rotating low-pressure turbine (LPT). The no. 4 roller bearing housing assembly is located within the TIC. The turbine intermediate case houses a turbine stator assembly which connects to the intermediate case using seven support rods. The turbine stator assembly is composed of 14 airfoils that direct HPT airflow to the LPT rotor.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-42
CS130_72_L2_TIC_General.fm
72 - Engine 72-52 Turbine Intermediate Case
Support Rod (7x)
Turbine Stator Assembly Airfoils
CS1_CS3_7251_003
No. 4 Roller Bearing Housing Assembly
Figure 21: Turbine Intermediate Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-43
CS130_72_L2_TIC_General.fm
72 - Engine 72-52 Turbine Intermediate Case
DETAILED COMPONENT INFORMATION TURBINE INTERMEDIATE CASE The turbine intermediate outer case holds the turbine intermediate inner case in a center position with seven rods. These rods run through the turbine stator vanes and provides support to the no.4 roller bearing housing assembly. It is oil damped in order to absorb rotor vibration, and supports the HPT drive shaft axially and radially. Heat shields are provided as thermal barriers to protect the no. 4 roller bearing housing assembly. Buffer air from the 4th stage provides sealing and cooling for the no. 4 bearing compartment, and prevents coking of engine oil. Lubrication and air cooling tubes are routed through the hollow turbine stator vanes to and from the no.4 roller bearing housing assembly. An outer heat shield provides a thermal barrier to the TICs outer case.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-44
CS130_72_L3_TIC_DetailedComponentInfo.fm
72 - Engine 72-52 Turbine Intermediate Case
Turbine Intermediate Case Rod (7x)
High-Pressure Turbine
Outer Heat Shield
Turbine Stator Vanes (14x)
Heat Shield Heat Shield Turbine Intermediate Inner Case
Heat Shield
No. 4 Roller Bearing Housing Assembly
High-Pressure Turbine Drive Shaft
CS1_CS3_7252_001
Low-Pressure Turbine Drive Shaft
Figure 22: Turbine Intermediate Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-45
CS130_72_L3_TIC_DetailedComponentInfo.fm
72 - Engine 72-53 Low-Pressure Turbine
72-53 LOW-PRESSURE TURBINE GENERAL DESCRIPTION The low-pressure turbine (LPT) drives the LPT shaft. The LPT case provides an attachment for the active clearance control (ACC) manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-46
CS130_72_L2_LPT_General.fm
72 - Engine 72-53 Low-Pressure Turbine
Active Clearance Control Manifold
Low-Pressure Turbine Case
CS1_CS3_7253_001
Low-Pressure Turbine Shaft
Figure 23: Low-Pressure Turbine Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-47
CS130_72_L2_LPT_General.fm
72 - Engine 72-53 Low-Pressure Turbine
DETAILED COMPONENT INFORMATION LOW-PRESSURE TURBINE The low-pressure turbine (LPT) consists of 3 rotor stages and 2 stages of stator vanes. The LPT drive shaft runs through the center of the engine. The LPT 2nd stage rotor hub is splined to the LPT drive shaft. The low-pressure turbine (LPT) 1st stage and 3rd stage rotor disks are bolted together on the LPT 2nd stage hub using tie rods. The LPT 1st stage rotor blades are hollow, while the 2nd and 3rd stage rotor blades are solid alloy. The LPT 2nd and 3rd stator vanes are segment arrangement type. Fan air is sprayed around the exterior of the LPT case from the active clearance control (ACC) cooling manifold. This reduces the LPT case diameter, reducing the turbine blades tip clearance, thus improving fuel efficiency. Insulation segments, located on the inside periphery of the LPT, protect the LPT case from heat.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-48
CS130_72_L3_LPT_DetailedComponentInfo.fm
72 - Engine 72-53 Low-Pressure Turbine
Insulation Segments
Low-Pressure Turbine Case
Active Clearance Control Cooling Manifold
Low-Pressure Turbine 1st Stage Rotor Blade
Low-Pressure Turbine 2nd Stage Stator Vane Low-Pressure Turbine 3rd Stage Rotor Blade
Low-Pressure Turbine 3rd Stage Stator Vane
Low-Pressure Turbine 2nd Stage Rotor Blade Tie Rod
Low-Pressure Turbine Drive Shaft
Low-Pressure Turbine 2nd Stage Rotor Hub
Low-Pressure Turbine 3rd Stage Rotor Disk
CS1_CS3_7253_002
Low-Pressure Turbine 1st Stage Rotor Disk
Figure 24: Low-Pressure Turbine Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-49
CS130_72_L3_LPT_DetailedComponentInfo.fm
72 - Engine 72-54 Turbine Exhaust Case
72-54 TURBINE EXHAUST CASE GENERAL DESCRIPTION The turbine exhaust case (TEC) is the aft most engine structural element. It is formed by the outer duct surface, the inner duct surface, eleven airfoil shaped struts, and the no.5 and no.6 bearing housings. The TEC is attached to the rear flange of the low-pressure turbine (LPT) at the front end, and to the exhaust system at the aft end. It forms a transition duct between the LPT and the exhaust case. Two ground handling bosses are provided at the 3 o’clock and 6 o'clock position. The TEC also houses the aft engine mount lugs at its top surface. Eight exhaust gas temperature (EGT) probes bosses are provided on the case outer surface. Oil supply and scavenge tubes use the hollow interior of two lower struts to reach no. 5 and no. 6 bearing housings.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-50
CS130_72_L2_TEC_General.fm
72 - Engine 72-54 Turbine Exhaust Case
Aft Engine Mount Lugs
Strut (11x)
Inner Duct Surface
Exhaust Gas Temperature Probe (8x)
No. 6 Bearing Housing Outer Duct Surface
No. 5 and No. 6 Bearing Oil Supply Tube
Ground Handling Boss (2x) No. 5 and No. 6 Bearing Oil Scavenge Tube
CS1_CS3_7254_001
No. 5 Bearing Housing
Figure 25: Turbine Exhaust Case Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-51
CS130_72_L2_TEC_General.fm
72 - Engine 72-54 Turbine Exhaust Case
DETAILED COMPONENT INFORMATION TURBINE EXHAUST CASE The turbine exhaust case (TEC) and struts attach to the no. 5 and no. 6 bearing housings. Heat shields are installed to protect both bearing housings from excessive hot exhaust gas flow. The no. 5 and no. 6 bearings are roller bearings that provide the radial support for the LPT shaft. Both bearings are oil-damped to absorb vibration. The no. 5 bearing rests against its forward seat and is properly kept separated from no.6 bearing by a shaft spacer. The no. 6 bearing retaining nut is installed on the low-pressure turbine (LPT) shaft to keep both bearing assemblies together. The three aft engine mount lugs carry part of the engine weight and loads to the pylon structure.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-52
CS130_72_L3_TEC_DetailedComponentInfo.fm
72 - Engine 72-54 Turbine Exhaust Case
Aft Engine Mount Lugs
Low-Pressure Turbine
Turbine Exhaust Case
Strut
Heat Shield
No. 6 Bearing Housing No. 6 Bearing
No. 6 Bearing Retaining Nut
No. 5 Bearing Housing Low-Pressure Turbine Shaft No. 5 Bearing Seat
No. 5 Bearing
Shaft Spacer
CS1_CS3_7254_002
Heat Shield
Figure 26: Turbine Exhaust Case Cross-Section Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-53
CS130_72_L3_TEC_DetailedComponentInfo.fm
72 - Engine 72-60 Main and Angle Gearboxes
72-60 MAIN AND ANGLE GEARBOXES The angle gearbox mechanically connects the high-pressure compressor (N2) and the main gearbox through the angle gearbox layshaft.
GENERAL DESCRIPTION MAIN GEARBOX The main gearbox (MGB) housing is a cast aluminum structure mounted to the engine core. Accessories are mounted on the forward and aft faces of the gearbox. The gearbox housing has the following components: •
Replaceable carbon seals
•
De-oiler
•
N2 crank pad port
•
Core drain plugs
•
Permanent magnet alternator generator (PMAG) drive
The MGB uses gears and shafts to drive accessories mounted to it. The angle gearbox layshaft drive is the rotational input to the MGB. The N2 speed is reduced before reaching the angle gearbox in order to have the MGB input drive in the proper rotational speed range. The gearbox aft face is axially supported by a rod link assembly that connects to the turbine intermediate case (TIC). A centering puck is installed on the gearbox top surface to help align the gearbox to the engine when it is installed. Two mounting links are used to mount the gearbox to the diffuser case. A heat shield protects the gearbox from engine core radiating heat. It is clipped at the front side and bolted to brackets at the aft end top surface of the gearbox.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-54
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
Heat Shield Left Mounting Link
Centering Puck
Air Turbine Starter Carbon Seal
Self Contained Deoiler N2 Cranking Port Pad Right Mounting Link Rod Link Assembly
Integrated Fuel Pump and Control Carbon Seal
Angle Gearbox Layshaft Drive Hydraulic Pump Carbon Seal
Deoiler Carbon Seal Permanent Magnet Alternator Generator (PMAG) Drive
Variable Frequency Generator Carbon Seal
CS1_CS3_7261_002
Core Drain Plugs
Lubrication and Scavenge Oil Pump Drive Pad
Figure 27: Main Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-55
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
MAIN GEARBOX MOUNTED COMPONENTS The following components are mounted on the main gearbox: •
Hydraulic pump
•
Lubrication and scavenge oil pump (LSOP)
•
Oil control module (OCM)
•
Variable frequency generator (VFG)
•
Permanent magnet alternator generator (PMAG)
•
Air turbine starter
•
Fuel oil manifold
•
Integrated fuel pump and control (IFPC)
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-56
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
Fuel/Oil Manifold
Integrated Fuel Pump and Control
Angle Gearbox Main Gear Box
Air Turbine Starter
Hydraulic Pump Lubrication and Scavenge Oil Pump
Oil Control Module Variable Frequency Generator
CS1_CS3_7261_001
Permanent Magnet Alternator Generator
Figure 28: Main Gearbox Mounted Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-57
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
ANGLE GEARBOX The angle gearbox mechanically connects the HPC (N2) and the main gearbox (MGB) through the layshaft. The angle gearbox is installed at the 8 o'clock position on the compressor intermediate case (CIC). The shaft is covered by two layshaft covers. The forward layshaft cover front end is secured to the angle gearbox, and the rear cover aft end is secured to the MGB. Both layshaft covers slide one over the other at the mating center area to allow for thermal growth during engine operation. When the engine is being started, the air starter drives the MGB, which is connected to the angle gearbox. The angle gearbox drives the HPC. When the engine is running, the HPC drives the MGB through the angle gearbox.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-58
CS130_72_L2_MAG_General.fm
72 - Engine 72-60 Main and Angle Gearboxes
Angle Gearbox
Layshaft
CS1_CS3_7260_001
Layshaft Covers
Figure 29: Angle Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-59
CS130_72_L2_MAG_General.fm
72 - Engine 72-00 Borescope Access
72-00 BORESCOPE ACCESS GENERAL DESCRIPTION
ENGINE LEFT SIDE BORESCOPE ACCESS
The borescope permits visual inspection of internal gaspath parts without the need for disassembly.
Table 1: Left Side Borescope Access Ports and Inspection Areas
A borescope probe is inserted in one of several access ports to inspect parts for damage, cracks, wear and missing material.
PORT
LOCATION
AP-LPC 3
9:30 position
LPC 2nd stage stator vanes and 3rd stage IBR blades.
AP-HPT 1
9 o’clock position
HPT 1st stage rotor blades, 2nd stage vanes, and 2nd stage rotor blades.
AP-LPT 1
8:30 position
LPT 1st stage rotor blades, LPT 1st stage vanes, and LPT 2nd stage rotor blades.
AP-LPT 2
8:30 position
LPT 2nd stage rotor blades, LPT 2nd stage vanes, and LPT 3rd stage rotor blades.
AP-DIFF 1
11 o’clock position
Combustion chamber inner and outer liners, bulkhead, fuel nozzles, HPT 1st stage rotor blades, and HPT 1st stage vanes.
Some damage to blades or stators can be blended in-situ using borescope blade blending. Consult the Aircraft Maintenance Publication (AMP) for airfoil damage and blend limitations. The following tables show the borescope access ports and inspection areas.
INSPECTION AREA
CS1_CS3_TB72_003
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-60
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
LPC 3
DIFF 1
HPT 1 LPT 1
CS1_CS3_7200_005
LPT 2
Figure 30: Engine Left Side Borescope Access Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-61
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
ENGINE RIGHT SIDE BORESCOPE ACCESS Table 2: Right Side Borescope Access Ports and Inspection Areas PORT
LOCATION
INSPECTION AREA
AP-LPC 1
3:30 position
LPC variable inlet stator vanes 1st stage integral bladed rotor (IBR) airfoils.
AP-LPC 2
4 o’clock position
LPC 1st stage stator vanes and 2nd stage IBR airfoils.
AP-HPC 1
4 o’clock position
HPC variable inlet guide vanes and 1st stage IBR airfoils.
AP-HPC 2
4 o’clock position
HPC 1st stage variable stator vanes (VSV) and 2nd stage IBR airfoils.
AP-HPC 3
3:30 position
HPC 2nd stage VSV and 3rd stage IBR airfoils.
AP-HPC 4
4 o’clock position
HPC 3rd stage VSV and 4th stage IBR airfoils.
AP-HPC 5
4.o’clock position
HPC4th stage stator and 5th stage IBR airfoils.
AP-HPC 6
4 o’clock position
HPC 5th stage stator and 6th stage IBR airfoils.
AP-HPC 7
2 o’clock position
HPC 6th stage stator and 7th stage IBR airfoils.
AP-HPC 8
2 o’clock position
HPC 7th stage stator and 8th stage rotor blade.
Igniter Plug Ports
3:30 and 4 o’clock positions
Combustion chamber inner and outer liner, bulkhead, fuel nozzles, HPT 1st stage rotor blades and vanes.
AP-TIC 1
1:30 position
HPT 2nd stage rotor blades. LPT 1st stage rotor blades. CS1_CS3_TB72_004
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-62
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
TIC 1
HPC 8 HPC 7
HPC 6
Igniter Upper
HPC 2 HPC 3 HPC 4 HPC 1
Igniter Lower
LPC 2
LPC 1
CS1_CS3_7200_006
HPC 5
Figure 31: Engine Right Side Borescope Access Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
72-63
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
BORESCOPE ACCESS - MAIN GEARBOX The gearbox also has two access ports to facilitate inspection. An N2 crank pad is used to turn the N2 rotor during the borescope procedure. BORESCOPE ACCESS - ANGLE GEARBOX The angle gearbox has one borescope access port for internal inspection.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-64
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
N2 Crank Pad Cover
Borescope Port
Borescope Port
Borescope Port Angle Gearbox
CS1_CS3_7200_007
Main Gearbox
Figure 32: Borescope Access - Gearbox and Angle Gearbox Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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72-65
CS130_72_L2_BorescopeAccess_General.fm
72 - Engine 72-00 Borescope Access
Page Intentionally Left Blank
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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72-66
CS130_72_ILB .fm
ATA 73 - Engine Fuel and Control
BD-500-1A10 BD-500-1A11
73 - Engine Fuel and Control
Table of Contents 73-11 Distribution....................................................................73-2 General Description ..........................................................73-2 Component Location ........................................................73-4 Component Information ....................................................73-6 Integrated Fuel Pump and Control...............................73-6 Fuel Filter .....................................................................73-8 Fuel Nozzles ..............................................................73-10 Detailed Description .......................................................73-12 Motive Flow Fuel........................................................73-12 Servo Fuel..................................................................73-12 Fuel Filter ...................................................................73-12 Metered Fuel..............................................................73-12 Fuel/Oil Heat Exchanger Operation ...........................73-12 73-20 Control System ...........................................................73-14 General Description ........................................................73-14 Electronic Engine Control ..........................................73-16 Fuel Control Description ............................................73-18 Prognostics and Health Management Unit ................73-20 Component Location ......................................................73-22 P2/T2 Probe...............................................................73-22 Burner Pressure Sensor ............................................73-22 Electronic Engine Control Unit ...................................73-22 Prognostics and Health Management Unit ................73-22 Data Storage Unit ......................................................73-22 Permanent Magnet Alternator Generator...................73-22 P2.5/T2.5 Probe.........................................................73-24 T3 Probe ....................................................................73-24 Component Information ..................................................73-26 Electronic Engine Control ..........................................73-26 Data Storage Unit ......................................................73-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Prognostics and Health Management unit ................ 73-26 Permanent Magnet Alternator Generator ................. 73-28 Detailed Component Information ................................... 73-30 Electronic Engine Control Power .............................. 73-30 Electronic Engine Control Fault Detection ................ 73-32 Prognostics and Health Management Unit................ 73-34 Data Storage Unit...................................................... 73-36 Controls and Indications ................................................ 73-38 Throttle Quadrant Assembly ..................................... 73-38 Engine Panel............................................................. 73-38 Engine Fire Panel...................................................... 73-38 Indications ................................................................. 73-40 Detailed Description ...................................................... 73-42 Aircraft and Electronic Engine Control Interface ....... 73-42 Engine Sensor Interface............................................ 73-44 Engine Operating Limits............................................ 73-46 Overspeed Protection, Shaft Shear Protection and Thrust Control Malfunction ................................. 73-48 Integrated Fuel Pump and Control Operation - Start ........................................... 73-50 Integrated Fuel Pump and Control Operation - Run ............................................ 73-52 Normal Engine Shutdown ........................................ 73-54 Emergency Shutdown ............................................... 73-56 Engine Power Management...................................... 73-58 Engine Power Settings and Modes ........................... 73-60 Automatic Power Reserve......................................... 73-62 N1 Synchronization ................................................... 73-64 Fault Reporting and Testing Interface....................... 73-66 Monitoring and Tests ...................................................... 73-68 CAS Messages ......................................................... 73-69
TECHNICAL TRAINING MANUAL
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73-i
CS130_M5_73_EngFuelCtlTOC.fm
73 - Engine Fuel and Control
Practical Aspects.............................................................73-72 Onboard Maintenance System ..................................73-72 73-30 Indicating ....................................................................73-76 General Description ........................................................73-76 Component Location ......................................................73-78 Controls and Indications .................................................73-80 Detailed Description .......................................................73-82 Monitoring and Tests.......................................................73-84 CAS Messages ..........................................................73-85
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-ii
CS130_M5_73_EngFuelCtlTOC.fm
73 - Engine Fuel and Control
List of Figures Figure 1: Fuel Distribution System General Description.......73-3 Figure 2: Fuel Distribution System Component Location .....73-5 Figure 3: Integrated Fuel Pump and Control ........................73-7 Figure 4: Fuel Filter ..............................................................73-9 Figure 5: Fuel Nozzles........................................................73-11 Figure 6: Fuel Distribution System Detailed Description ....73-13 Figure 7: Engine Control System General Description.......73-15 Figure 8: Electronic Engine Control....................................73-17 Figure 9: Fuel Control.........................................................73-19 Figure 10: Prognostics and Health Management Unit..........73-21 Figure 11: Engine Control System Component Location .....73-23 Figure 12: P2.5/T2.5 Probe and T3 Probe Location.............73-25 Figure 13: Electronic Engine Control, Data Storage Unit, and Prognostics and Health Management Unit..........73-27 Figure 14: Permanent Magnet Alternator Generator ............73-29 Figure 15: Electronic Engine Control Power.........................73-31 Figure 16: Electronic Engine Control Fault Detection...........73-33 Figure 17: Prognostics and Health Management Unit..........73-35 Figure 18: Data Storage Unit................................................73-37 Figure 19: Fuel Control System Controls ............................73-39 Figure 20: Fuel Control System Indications.........................73-41 Figure 21: Aircraft and Electronic Engine Control Interface..................................................73-43 Figure 22: Engine Sensor Interface......................................73-45 Figure 23: Electronic Engine Control Engine Operating Limits ..................................................73-47 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 24: Overspeed, Shaft Shear Protection, and Thrust Control Malfunction........................... 73-49 Figure 25: Integrated Fuel Pump and Control Operation - Engine Start ........................ 73-51 Figure 26: Integrated Fuel Pump and Control Operation - Engine Run ......................... 73-53 Figure 27: Normal Engine Shutdown ................................... 73-55 Figure 28: Emergency Shutdown........................................ 73-57 Figure 29: Engine Bleed Management and Idle Speed ....... 73-59 Figure 30: Engine Power Settings and Modes..................... 73-61 Figure 31: Automatic Power Reserve .................................. 73-63 Figure 32: N1 Synchronization............................................. 73-65 Figure 33: Electronic Engine Control Fault Reporting and Testing Interface ......................... 73-67 Figure 34: OMS EEC System Test ...................................... 73-73 Figure 35: Power Assurance Test........................................ 73-75 Figure 36: Fuel Control Indicating System........................... 73-77 Figure 37: Fuel Control Indicating System Component Location........................................... 73-79 Figure 38: Fuel Control Indications ...................................... 73-81 Figure 39: Fuel Control Indicating Detailed Description....... 73-83
TECHNICAL TRAINING MANUAL
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73-iii
CS130_M5_73_EngFuelCtlLOF.fm
ENGINE FUEL AND CONTROL - CHAPTER BREAKDOWN ENGINE FUEL AND CONTROL CHAPTER BREAKDOWN
Distribution
1
Control System
2
Indicating
3
73 - Engine Fuel and Control 73-11 Distribution
73-11 DISTRIBUTION GENERAL DESCRIPTION The engine fuel distribution system supplies fuel to the engine for combustion and for air system actuator operation. The aircraft fuel system supplies fuel to the integrated fuel pump and control (IFPC). A fuel oil manifold provides mounting for the IFPC, fuel flow transmitter and the fuel filter. The manifold simplifies maintenance by minimizing the number of external fuel lines needed to connect the components. The IFPC has integral pumps driven by the engine gearbox: •
Centrifugal, low-pressure boost pump
•
High-pressure main gear pump to provide pressurized fuel
•
Motive flow gear pump
The filtered fuel then flows to the main pump where it is pressurized. Some of the pressurized fuel is supplied to the air system actuators for engine cooling and stability control. The main pump also supplies the fuel to the fuel metering valve (FMV). The FMV is controlled by the electronic engine control (EEC) to provide the correct fuel for a given thrust setting. The metered fuel flow is measured by the fuel flow meter. The fuel is then sent to the flow divider. The flow divider valve (FDV) allows fuel to flow to the primary fuel nozzles for engine start. As the engine accelerates to idle, the flow divider valve provides flow to the secondary fuel nozzles.
Fuel from the aircraft is supplied to the IFPC boost pump under pressure. Should the aircraft fuel system pumps fail, the boost pump is capable of suctioning fuel from the aircraft fuel tanks. The boost pump provides fuel at low-pressure to the motive flow pump and to a fuel/oil heat exchanger (FOHX). The motive flow pump provides fuel to operate the aircraft fuel system ejector pumps. The FOHX heats engine fuel to prevent icing and cools the engine oil for proper engine bearing seal operation. The heated fuel then passes through the fuel filter, which is installed in a housing attached to the fuel manifold.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-2
CS130_73_L2_Distribution_General.fm
73 - Engine Fuel and Control 73-11 Distribution
INTEGRATED FUEL PUMP AND CONTROL PRESSURE-REGULATING VALVE Fuel From Main Tank
Boost Pump MAIN PUMP
LEGEND
ACC EEC FDV FMV FOHX HPC LPC TM
Engine Oil Fuel Metered Fuel Motive Flow Pressurized Fuel Pump Discharge Servo Fuel Servo Fuel Return Active Clearance Control Electronic Engine Control Flow Divider Valve Fuel Metering Valve Fuel/Oil Heat Exchanger High-Pressure Compressor Low-Pressure Compressor Torque Motor
FMV
MOTIVE FLOW PUMP
FDV
FUEL/OIL MANIFOLD To Primary Manifold To Secondary Manifold
To Motive Flow System
To Engine Oil
FOHX
TURBINE HPC LPC ACC STATOR STATOR AIR VANES VANE VALVE FUEL FILTER
2.5 BLEED VALVE
EEC
CS1_CS3_7311_001
From Engine Oil
FUEL FLOW METER
Figure 1: Fuel Distribution System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-3
CS130_73_L2_Distribution_General.fm
73 - Engine Fuel and Control 73-11 Distribution
COMPONENT LOCATION
Fuel/Oil Manifold
The engine fuel distribution system consists of the following components:
The fuel/oil manifold is located on the front face of the gearbox on the left side of the engine.
•
Integrated fuel pump and control
•
Fuel/oil heat exchanger
•
Fuel/oil heat exchanger bypass valve
•
Fuel filter
•
Fuel nozzles
•
Fuel supply manifolds
•
Fuel/oil manifold
•
Fuel flow meter
Fuel Flow Meter The fuel flow meter is attached to the fuel oil manifold on the front face of the main gearbox at the 9 o’clock position.
Integrated Fuel Pump and Control The integrated fuel pump and control (IFPC) is mounted on the fuel/oil manifold on the front face of the main gearbox at the 8 o’clock position. Fuel/Oil Heat Exchanger and Fuel/Oil Heat Exchanger Bypass Valve The fuel/oil heat exchanger (FOHX) and fuel/oil heat exchanger bypass valve are located at the 12 o’clock position on the compressor intermediate case. Fuel Filter The fuel filter is located below the IFPC on the fuel/oil manifold. Fuel Nozzles and Primary and Secondary Fuel Manifolds The fuel nozzles and primary and secondary fuel manifolds are located on the circumference of the combustor.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-4
CS130_73_L2_Distribution_ComponentLocation.fm
73 - Engine Fuel and Control 73-11 Distribution
Fuel/Oil Heat Exchanger Bypass Valve (FOHXBV)
Fuel/Oil Heat Exchanger (FOHX) Fuel Supply Manifolds
Fuel Flow Meter
Fuel Nozzles
Fuel/Oil Manifold
Fuel Filter
CS1_CS3_7311_002
Integrated Fuel Pump and Control (IFPC)
Figure 2: Fuel Distribution System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-5
CS130_73_L2_Distribution_ComponentLocation.fm
73 - Engine Fuel and Control 73-11 Distribution
COMPONENT INFORMATION INTEGRATED FUEL PUMP AND CONTROL The IFPC supplies metered fuel based on inputs from the EEC in response to changes in thrust demand and environmental conditions. The IFPC performs the following major functions: •
Supply motive fuel for aircraft fuel system
•
Meter fuel for combustion
•
Pressurize fuel supply
•
Fuel shutoff for engine shutdown
•
Control primary and secondary flow distribution to fuel nozzles
•
Supply pressurized servo fuel for air system actuators
•
Bypass excess fuel flow
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-6
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
MOTIVE FLOW
METERED FUEL FLOW
FUEL PRESSURE REGULATION
FUEL SHUTOFF
PRIMARY AND SECONDARY MANIFOLD PRESSURIZATION
INTEGRATED FUEL PUMP AND CONTROL
SERVO FUEL
CS1_CS3_7311_003
EXCESS FUEL BYPASS
Figure 3: Integrated Fuel Pump and Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-7
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
FUEL FILTER The main fuel filter element is used to remove contaminants from the fuel sent from the IFPC. The canister type disposable fuel filter is installed in a housing on the fuel oil manifold below the IFPC. The fuel filter bypass valve opens when the filter clogs and the differential pressure exceeds 25 psid. The filter housing is equipped with a drain plug for draining the housing prior to removing the filter.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-8
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
Filter Housing Drain
Fuel Filter Bypass Valve
CS1_CS3_7311_004
Fuel Filter Housing
Figure 4: Fuel Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-9
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
FUEL NOZZLES There are 16 fuel nozzles, 10 duplex, and 6 simplex. They are arranged in a pattern of 4 duplex, 4 simplex, 6 duplex, and 2 simplex nozzles. The IFPC and fuel/oil manifold send fuel to the primary and secondary fuel nozzle manifolds that supply the fuel nozzles. The duplex nozzles contain both primary fuel nozzles and secondary fuel nozzles. The simplex nozzles only contain secondary fuel nozzles. All of the fuel nozzles contain check valves to prevent fuel in the primary and secondary fuel nozzle manifolds from draining into the engine following an engine shutdown. These valves in the primary and secondary circuits of the fuel nozzles shut when the fuel system pressure drops below a minimum threshold during shutdown conditions, preventing fuel from draining into the combustor. During an engine start, the IFPC delivers fuel to the primary manifold to provide combustion for initial spool up of the engine. During engine acceleration to ground idle, the FDV in the IFPC distributes fuel flow to both the primary and secondary nozzles. An alignment pin on each nozzle prevents incorrect installation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-10
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
Primary Flow Secondary Flow
Simplex Nozzle
Secondary Manifold
Duplex Nozzle Alignment Pin
Primary Manifold
DUPLEX NOZZLE
Secondary Flow FUEL NOZZLES (AFT LOOKING FORWARD)
FUEL NOZZLES ON COMBUSTION CHAMBER
SIMPLEX NOZZLE
CS1_CS3_7311_005
Alignment Pin
Figure 5: Fuel Nozzles Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-11
CS130_73_L2_Distribution_ComponentInformation.fm
73 - Engine Fuel and Control 73-11 Distribution
DETAILED DESCRIPTION MOTIVE FLOW FUEL Motive flow for the aircraft fuel system is provided by a motive flow gear pump downstream of the centrifugal boost pump exit. The inlet of the motive flow pump is protected from fuel contamination by a screen with a bypass valve. The outlet of the motive flow pump is protected from overpressure by a high-pressure relief valve. SERVO FUEL The servo pressure regulator adjusts the pressure to a suitable level for actuator operation. Servo fuel is supplied to the turbine active clearance control (ACC) valve, high-pressure compressor (HPC) stator vane actuator, low-pressure compressor (LPC) stator vane actuator, and the 2.5 bleed valve actuator. The actuators are bi-directional fully modulated linear actuators used to optimize overall engine performance and to prevent engine compressor surges. The servo fuel is then returned from the actuators upstream of the fuel filter before being recirculated in the IFPC.
The metered fuel output of the FMV then flows to the fuel flow meter as long as the windmill bypass valve (WBV) remains closed. The fuel flow meter sends a flow signal to the EEC for fuel flow indication. Fuel is supplied to the minimum pressure-shutoff valve (MPSOV) and flow divider valve (FDV) for distribution to the fuel nozzles. When the engine is shutdown, the WBV opens and connects fuel from upstream of the minimum pressure-shutoff valve (MPSOV) with the fuel filter inlet. The MPSOV closes and fuel is shut off from the engine. FUEL/OIL HEAT EXCHANGER OPERATION The fuel/oil heat exchanger maintains the fuel temperature above 0°C (32°F) at the fuel filter inlet during steady-state operation by transferring engine oil heat to the fuel. The EEC controls the fuel/oil heat exchanger bypass valve (FOHXBV) based on the fuel temperature sensor signal. As the temperature decreases, the EEC opens the FOHXBV to supply more hot oil to warm the fuel. If the fuel side of the heat exchanger is blocked, a bypass valve opens to divert fuel around the heat exchanger.
FUEL FILTER The fuel filter differential pressure sensor measures the pressure drop across the fuel filter. If the fuel filter is completely clogged, a bypass valve opens to divert fuel around the fuel filter. The EEC monitors the fuel filter and provides an L(R) ENG FUEL FILTER advisory message when the fuel filter clogs. An accompanying INFO message indicates a filter clogging or an actual bypass condition. METERED FUEL The pressure-regulating valve (PRV) modulates to provide a constant fuel flow to the FMV regardless of the FMV position. The EEC computes the required metered fuel using the fuel temperature sensor for compensation. In the case of a fuel temperature sensor failure, the EEC uses a default fuel temperature of 75°C (167°F). Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-12
CS130_73_L3_Distribution_DetailedDescription.fm
73 - Engine Fuel and Control 73-11 Distribution
INTEGRATED FUEL PUMP AND CONTROL LEGEND
FTS HPC LPC MPSOV T/M WBV
¨3
Fuel Flow Differential Pressure Sensor
PRESSURE REGULATING VALVE Boost Pump WBV
MAIN PUMP FTS FMV MOTIVE FLOW PUMP
Servo Pressure Regulator MPSOV
FDV
FUEL/OIL MANIFOLD
To Primary Manifold To Secondary Manifold
Screen To Motive Flow
FUEL FLOW METER T/M
¨3
FOHX
TURBINE HPC LPC ACC STATOR STATOR AIR VANES VANE VALVE
From Engine Oil FOHXBV To Engine Oil
FUEL FILTER
2.5 BLEED VALVE
EEC
CS1_CS3_7311_007
ACC EEC FDV FMV FOHX FOHXBV
Engine Oil Fuel From Fuel Main Tank Metered Fuel Motive Flow Pressurized Fuel Pump Discharge Servo Fuel Servo Fuel Return Active Clearance Control Electronic Engine Control Flow Divider Valve Fuel Metering Valve Fuel/Oil Heat Exchanger Fuel/Oil Heat Exchanger Bypass Valve Fuel Temperature Sensor High-Pressure Compressor Low-Pressure Compressor Minimum Pressure-Shutoff Valve Torque Motor Windmill Bypass Valve Bypass Valve
Figure 6: Fuel Distribution System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-13
CS130_73_L3_Distribution_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
73-20 CONTROL SYSTEM GENERAL DESCRIPTION The engine is controlled by a full authority digital engine control (FADEC). A dual-channel electronic engine control (EEC) controls the engine thrust, using N1 as a reference, in accordance with the position of the throttle quadrant assembly (TQA) thrust lever, environmental conditions, and the amount of bleed air being extracted from the engine. This is done by the EEC using the data received from both aircraft systems and engine sensors, together with the engine trim data inputs from the data storage unit (DSU).
A gearbox-driven permanent magnet alternator generator (PMAG) provides electrical power to the EEC during normal operation. The EEC channel A and channel B, each have their own processor, power supply, program memory, and inputs and outputs. Either EEC channel is capable of controlling all of the critical engine functions: •
Engine fuel control
•
Engine starting control
The FADEC consists of three main components:
•
Engine stability
•
Electronic engine control (EEC)
•
Engine limit control
•
Prognostics and health management unit (PHMU)
•
Data storage unit (DSU)
During normal operation, one channel is in control and the other channel is in standby. Both channels calculate and provide the same commands, however, only the outputs from the channel in control are active. In case of an output failure on the channel in control, the equivalent output from the standby channel is used and controlled by the standby channel.
The data storage unit (DSU) is installed on the EEC and used for storage of engine configuration and identification data, including information related to the engine’s specific serial number. The prognostics and health management unit (PHMU) is a single channel line replaceable unit (LRU) that monitors engine parameters to provide the engine health status, vibration, auxiliary oil pressure sensing, and oil debris monitoring. The software in the EEC provides the logic required for engine control, fault detection and accommodation. The engine indication and crew alerting system (EICAS) display parameters, and engine diagnostics and maintenance data to the onboard maintenance system (OMS). The EEC also provides health monitoring information from the PHMU to the health management unit (HMU). All of the FADEC components have fieldloadable software.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-14
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
Aircraft Systems
DMC 1 and DMC 2
EICAS SYN OMS
THROTTLE QUADRANT ASSEMBLY
HMU ELECTRONIC ENGINE CONTROL
DSU
Channel A Engine Fuel Control Engine Starting Control Engine Stability Engine Limit Protection Channel B
ENGINE SENSORS DMC DSU
Data Concentrator Unit Module Cabinet Data Storage Unit ARINC 429 CAN BUS
PERMANENT MAGNET ALTERNATOR GENERATOR
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
NOTE Left engine shown. Right engine similar.
CS1_CS3_7321_012
LEGEND
Figure 7: Engine Control System General Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-15
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
ELECTRONIC ENGINE CONTROL
Cowl Anti-Ice
The EEC channels are normally powered by the PMAG when engines are running but are powered by ESSENTIAL DC BUSES when the engines and the PMAG fails or the engine is shutdown. The left EEC channel A is powered by DC ESS BUS 3 and channel B is powered by DC ESS BUS 1. The right EEC channel A is powered by DC ESS BUS 3, and channel B is powered by DC ESS BUS 2.
The cowl anti-ice (CAI) valves are controlled by the EEC to provide anti-icing for the inlet cowl. They are also opened to improve HPC stability during engine start and at low engine power settings.
Fuel Control The EEC schedules engine thrust in response to thrust lever angle from the throttle quadrant assembly (TQA), aircraft inputs from the integrated air management system (IAMS), flight management system (FMS) and engine sensor inputs. Engine N1 and N2 limits are protected at all times and the exhaust gas temperature (EGT) is limited at start. The main input for fuel flow scheduling is thrust lever angle from the TQA.
Air Systems The engine interfaces with the IAMS to provide the correct bleed demand based on air conditioning and wing anti-ice demands. Thrust Reverser The EEC controls the deploy and stow functions of the thrust reverser actuation system (TRAS). The EEC control is based on thrust lever position from the TQA when the aircraft is on the ground.
Additional parameters the EEC uses to calculate fuel flow are ambient pressure (Pambient), inlet air temperature (T2), and pressure (P2), burner pressure (Pb), exhaust gas temperature (EGT), N1 speed and N2 speed. Starting and Ignition System The control system provides automatic starting sequencing of the starter air valve for the air turbine starter, igniters, and hydraulic pump depressurization during the starting phase. Engine Airflow Control System The EEC controls bleed valves, and the low-pressure compressor (LPC) and high-pressure compressor (HPC) stator vane system for engine stability and operability. The EEC also controls the turbine active clearance control systems for improved engine efficiency.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-16
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
THROTTLE QUADRANT ASSEMBLY
ELECTRONIC ENGINE CONTROL DC ESS BUS 3
FUEL CONTROL DSU STARTING AND IGNITION SYSTEM
Channel A ENGINE AIRFLOW CONTROL
PERMANENT MAGNET ALTERNATOR GENERATOR DC ESS BUS 1
COWL ANTI ICE Channel B AIR SYSTEMS
THRUST REVERSER ENGINE SENSORS DSU
Data Storage Unit ARINC 429 CAN BUS
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
NOTE Left engine shown. Right engine similar.
CS1_CS3_7321_032
LEGEND
Figure 8: Electronic Engine Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-17
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
FUEL CONTROL DESCRIPTION Each engine is controlled by a thrust lever on the TQA. The thrust lever controls both forward and reverse thrust. The ENG switches provide engine run and shutdown commands for normal operation. The ENGINE panel has a START switch to initiate the start sequence in manual mode.
The fuel scheduling and limit protection is based on the following sensors: •
N1 speed
•
N2 speed
•
Ambient air pressure (Pamb)
The EEC uses information from the air data system for engine control. The air data is validated against data received from the engine sensors P2T2 and Pamb. Engine sensor data is used in the event the air data is not available, however, engine performance is degraded.
•
Inlet air temperature (T2)
•
Inlet pressure (P2)
•
Burner pressure (Pb)
The flight management system (FMS) provides thrust requirements based on the flight plan information entered. The FMS also provides an interface to manually control the thrust reference.
•
Exhaust gas temperature (EGT)
If a fire is detected, the engine is automatically shut down when the L(R) FIRE PBA is pressed.
The integrated air management system supplies the bleed air requirements from the aircraft systems. The thrust requirement is adjusted to ensure that the minimum bleed air requirements are always met. When the thrust requirements are set, the autothrottle can be engaged to control the thrust levers automatically, based on the FMS flight plan inputs. The EEC provides fuel scheduling and engine start commands to the integrated fuel pump and control (IFPC). The IFPC provides metered fuel for combustion and servo fuel for control of the engine air systems.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Engine rotor speeds are calculated by the EEC using the N1 and N2 signals. The EEC schedules fuel according to the calculated N1 target and limits fuel if N1 or N2 limits are reached. If an overspeed is detected on N1 or N2, the engine is shutdown. Burner pressure (Pb) is sent to the EEC to schedule fuel, detect stalls, and hung starts. The exhaust gas temperature (EGT) is monitored by the EEC. During engine start, the maximum EGT is limited. The EGT is used for temperature exceedance determination, light off detection, hot start abort and engine health monitoring. The EEC limits fuel flow when redline limits are reached for N1, N2 or burner pressure. In addition to redline limiting, if the high or low rotor speed exceeds the overspeed limit of 105% N1 or 105% N2, the EEC will command an overspeed solenoid in the integrated fuel pump and control unit (IFPC) to cut engine fuel flow.
TECHNICAL TRAINING MANUAL
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73-18
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
L ENG
F IR E BTL 1
BTL 2
AVAIL
AVAIL
FMS AIR DATA
L ENGINE FIRE PANEL AUTO THROTTLE ENGINE CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
DMC 1 and DMC 2
INTEGRATED AIR MANAGEMENT SYSTEM
ENGINE PANEL
MAX
ELECTRONIC ENGINE CONTROL IDLE
INTEGRATED FUEL PUMP AND CONTROL
IDLE
L ENG
ON
FF FF OFF
FF FF OFF
Servo Fuel
N1 N2 Pamb T2 P2 Pb EGT
R ENG
ON
Metered Fuel For Combustion
THROTTLE QUADRANT ASSEMBLY LEGEND DMC
NOTE
Data Concentrator Unit Module Cabinet ARINC 429
Left engine shown. Right engine similar.
CS1_CS3_7321_018
MAX
Figure 9: Fuel Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-19
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
PROGNOSTICS AND HEALTH MANAGEMENT UNIT The prognostics and health management unit (PHMU) provides information about the health of the engine to help optimize maintenance operations. The PHMU is a single channel system with health monitoring functions using data from the EEC and engine sensors. The PHMU also performs the following functions: •
Engine vibration monitoring
•
Fan trim balance
•
Oil debris monitoring
•
Fan drive gear system (FDGS) auxiliary oil pressure sensor monitoring
The PHMU is powered from DC BUS 1. The PHMU communicates with the EEC through a CAN BUS. The engine health monitoring information is supplied to the HMU where is can be downloaded for analysis. The engine vibration monitoring helps locate sources of vibration and provides information for fan trim balance. The fan trim balance is carried out through the onboard maintenance system (OMS). The PHMU monitors the fan drive gear system (FDGS) to ensure that it receives a constant supply of oil under all flight conditions.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-20
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
EICAS DMC 1 and DMC 2
OMS HMU
DC BUS 1
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
ENGINE SENSORS
Engine Data
ELECTRONIC ENGINE CONTROL
P2/T2 Sensors Engine Vibration Oil Debris Monitoring Fan Drive Gear System Auxiliary Oil Pressure Sensor LEGEND Data Concentrator Unit Module Cabinet ARINC 429 CAN BUS
CS1_CS3_7320_004
DMC
Figure 10: Prognostics and Health Management Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-21
CS130_73_L2_ControlSystem_General.fm
73 - Engine Fuel and Control 73-20 Control System
COMPONENT LOCATION
PERMANENT MAGNET ALTERNATOR GENERATOR
Components of the engine control system include:
The permanent magnet alternator generator (PMAG) is located on the main gearbox.
•
P2/T2 probe
•
Burner pressure (Pb) Sensor
•
Electronic engine control (EEC)
•
Prognostics and health management unit (PHMU)
•
Data storage unit (DSU)
•
Permanent magnet alternator generator (PMAG)
•
P2.5/T2.5 probe (Refer to Figure 12)
•
T3 probe (Refer to Figure 12)
P2/T2 PROBE The P2/T2 probe is mounted to the inlet cowl at the 11 o’clock position. BURNER PRESSURE SENSOR The Pb sensor is attached to the compressor intermediate case (CIC) fire seal at the 10 o’clock position. ELECTRONIC ENGINE CONTROL UNIT The EEC is located on the fan case at the 9 o’clock position. PROGNOSTICS AND HEALTH MANAGEMENT UNIT The PHMU is located beside the EEC on the engine fan case at the 9 o’clock position. DATA STORAGE UNIT The data storage unit (DSU) is located on the EEC.
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73-22
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
A
B BURNER PRESSURE (Pb) SENSOR
B C
A P2/T2 PROBE
Data Storage Unit (DSU)
Prognostic and Health Management Unit (PHMU)
D
C
PERMANENT MAGNET ALTERNATOR GENERATOR (PMAG)
CS1_CS3_7320_001
Electronic Engine Control (EEC)
D
Figure 11: Engine Control System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-23
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
P2.5/T2.5 PROBE The P2.5/T2.5 probe is located on the CIC at the 1 o’clock position. T3 PROBE The T3 probe is located on the diffuser case, forward of the fuel nozzles, at approximately the 1 o’clock position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-24
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
A B
B T3 PROBE
CS1_CS3_7320_003
A P2.5/T2.5 PROBE
Figure 12: P2.5/T2.5 Probe and T3 Probe Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-25
CS130_73_L2_ControlSystem_ComponentLocation.fm
73 - Engine Fuel and Control 73-20 Control System
COMPONENT INFORMATION ELECTRONIC ENGINE CONTROL
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
The EEC electronic components are installed in a cast aluminum housing that is cooled by natural convection using fan compartment air.
The PHMU electronic components are installed in a cast aluminum housing that is cooled by natural convection using fan compartment air.
The EEC is installed on the fan case using four vibration isolators. The EEC has eight electrical connectors that allow it to interface with the aircraft systems and engine sensors, a data storage unit (DSU), and two pneumatic pressure ports.
The PHMU is installed on the fan case using four vibration isolators. The PHMU has two electrical connectors that allow it to interface with the EEC and engine sensors.
The EEC has an internal ambient pressure sensor that receives ambient pressure through a port on the lower right side of the EEC. The EEC also receives inlet pressure (P2) through the P2 port. DATA STORAGE UNIT The DSU is used for storage of engine configuration and identification data, engine historical and exceedance information, thrust ratings data, and fan trim balance data. If an EEC or prognostics and health management unit (PHMU) is replaced, the data stored in the DSU is supplied to the new unit. The DSU is mounted on the EEC A channel housing. The DSU is attached to the engine fan case by a lanyard and remains with the engine if the EEC is replaced.
CAUTION The DSU must remain with the engine when the EEC is removed. An incorrectly installed DSU can affect engine operation.
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73-26
CS130_73_L2_ControlSystem_ComponentInformation.fm
73 - Engine Fuel and Control 73-20 Control System
Electrical Connectors (8x)
Data Storage Unit Plug (DSU)
Vibration Isolators (4x)
P2 Sense Port
Electronic Engine Control (EEC)
Ambient Pressure Port
Electrical Connectors
Vibration Isolators (4x)
CS1_CS3_7321_006
Prognostics And Health Management Unit (PHMU)
Figure 13: Electronic Engine Control, Data Storage Unit, and Prognostics and Health Management Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-27
CS130_73_L2_ControlSystem_ComponentInformation.fm
73 - Engine Fuel and Control 73-20 Control System
PERMANENT MAGNET ALTERNATOR GENERATOR The permanent magnet alternator generator (PMAG) consists of a stator and a rotor. The stator and electrical windings are attached to the gearbox with captive bolts. The rotor remains attached to the gearbox shaft when the stator is removed. The PMAG stator has dual-channel EEC AC power windings, dual-channel N2 windings and a single generator winding for AC power to the fly-by-wire (FBW) controller.
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73-28
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73 - Engine Fuel and Control 73-20 Control System
Permanent Magnet Alternator Generator Stator
LEGEND Electronic Engine Control Channel A Power Electronic Engine Control Channel B Power Fly-By-Wire Power Speed (N2) Coils
CS1_CS3_7321_002
Permanent Magnet Alternator Generator Rotor
Figure 14: Permanent Magnet Alternator Generator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-29
CS130_73_L2_ControlSystem_ComponentInformation.fm
73 - Engine Fuel and Control 73-20 Control System
DETAILED COMPONENT INFORMATION ELECTRONIC ENGINE CONTROL POWER During normal operating conditions with the engine operating above 45% N2 speed, the EEC is powered by the permanent magnet alternator generator (PMAG). The PMAG provides a separate output for each EEC channel. When the engine is not running, both EEC channels are powered by the aircraft electrical system. Channel A of each EEC is powered by DC ESS BUS 3. Channel B of the left EEC is powered by DC ESS BUS 1. Channel B of the right EEC is powered by DC ESS BUS 2. When the engine is at or above minimum idle, a failure of one PMAG output results in the corresponding EEC channel switch its aircraft electrical power source. The other channel of the EEC continues operation using power from the PMAG.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-30
CS130_73_L3_ControlSystem_DetailedComponentInfo.fm
73 - Engine Fuel and Control 73-20 Control System
L PMAG
R PMAG
N2 is Greater Than 45% 34 VAC
N2 is Greater Than 45% 34 VAC
LEFT CBP
RIGHT CBP
CDC 1-5-1
CDC 2-5-1
L FADEC B DC ESS BUS 3
10
L FADEC A
DC ESS BUS 1
SSPC 10A
L FADEC B DC ESS BUS 3
10
R FADEC A
Logic: Always On
R EEC
Channel B
Channel A
Channel B
CS1_CS3_7321_014
LEGEND CBP CDC EEC PMAG
SSPC 10A
Logic: Always On
L EEC
Channel A
DC ESS BUS 2
Circuit Breaker Panel Control and Distribution Cabinet Electronic Engine Control Permanent Magnet Alternator Generator
Figure 15: Electronic Engine Control Power Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
ELECTRONIC ENGINE CONTROL FAULT DETECTION The EEC is designed to be fail-safe by detecting and accommodating component failures within the system. This is accomplished by the fault accommodation logic in the software. The main goal of the fault accommodation logic is to make the system operational upon the detection of all single electrical faults when the system is dispatched with two channels operational. The secondary goal is to maintain as much engine functionality as possible when multiple faults exist.
In the active-standby control scheme, the system gives active control to the healthiest channel. When the engine is shutdown, control of the engine switches to the other channel for the next engine start. This aids in the detection of potential hidden failures and provides equal output driver usage in the EEC. The channel switchover function is tested on each EEC power application.
Failures detectable by the control system are classified by their effect on the ability of the channel to control the engine. Many input failures have redundant backup sources and therefore do not require a channel switchover. The system is designed such that the minimum number of channel switchovers will occur due to faults. The fault accommodation is based on: •
If there are no detected faults in either channel, use average of the two channels
•
If an input fault or faults exist, use data from the other channel
•
If an output fault or faults exist, switch to other channel output
•
In the case of a dual failure, the EEC fails to a designated safe mode
The active-standby control scheme is used until the first failure is detected. At that point, control switches over to the active driver control scheme. The channel switchover logic has active driver and active-standby control schemes. The active driver control scheme allows either channel to control any of the output drivers independently, regardless of which channel is in control.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
COMPONENT
ELECTRONIC ENGINE CONTROL
COMPONENT
Channel A
Channel A
Channel A
Channel A
Channel B
Channel B
Channel B
Channel B
NORMAL OPERATION
INPUT SWITCH
Both channels operating normally
Output from Channel A used
Channel A in control
Feedback from Channel B channels used
ELECTRONIC ENGINE CONTROL
COMPONENT
ELECTRONIC ENGINE CONTROL
COMPONENT
Channel A
Channel A
Channel A
Channel A
Channel B
Channel B
Channel B
Channel B
OUTPUT DRIVER SWITCH
CHANNEL SWITCH
Output from Channel B used
Channel A fault detected
Feedback from both channels used
Channel B in control
LEGEND Control Fault Feedback Standby
NOTE Channel A in control.
CS1_CS3_7321_027
ELECTRONIC ENGINE CONTROL
Figure 16: Electronic Engine Control Fault Detection Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
PROGNOSTICS AND HEALTH MANAGEMENT UNIT The prognostics and health management unit (PHMU) performs engine health monitoring functions based on data provided by the EEC. The PHMU also performs the following functions: •
Engine vibration monitoring
•
Oil debris monitoring
•
Fan drive gear system (FDGS) oil monitoring
The PHMU provides engine health monitoring by comparing streaming engine data from engine sensors to stored models of engine gas path and subsystem parameters. The P2.5/T2.5 sensor provides P2.5 pressure to the PHMU. The other engine sensors provide information through the EEC. The information is processed by the PHMU and the appropriate output is sent to the EEC. The EEC then sends this information to the onboard maintenance system (OMS) and the health management unit (HMU). The PHMU receives signals from engine vibration sensors to assess the health of rotating and mechanical components. The vibration signals are compared against the engine speed signals supplied by the EEC to determine the location of the vibration. The processed vibration signals are sent to EICAS and also provide additional vibration information for fan balancing. The oil debris monitoring system checks for metal particles in the oil system and provides an early indication of engine wear. The PHMU supplies the information to the EEC as part of the data report supplied to the HMU. The PHMU also monitors the FDGS oil supply to ensure that a continuous supply of oil is being provided.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
CDC 1-10-8 L PHMU DC BUS 1
SSPC 5A
FWD VIBRATION SENSOR
P2.5/T2.5 PROBE
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
OIL DEBRIS MONITOR SENSOR
AFT VIBRATION SENSOR
AUXILIARY OIL PRESSURE SENSOR
Logic: Always On
EEC N1 Speed N2 Speed Fan Speed Engine Sensors
CDC EEC
Control and Distribution Cabinet Electronic Engine Control CAN BUS
NOTE Left engine shown. Right engine similar.
CS1_CS3_7321_015
LEGEND
Figure 17: Prognostics and Health Management Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
DATA STORAGE UNIT
Engine Thrust Rating
The digital storage unit (DSU) connects directly to EEC channel A. It communicates with both channels through a serial data bus. Data is stored in the DSU memory, EEC channel A, and EEC channel B.
The EEC uses the engine nameplate identification and aircraft type on the DSU to select which set of thrust ratings data to transfer to processor memory. During synchronization, the thrust ratings data corresponding to the synchronized engine nameplate identification and aircraft type are transferred to the EEC processor memory.
The DSU is capable of storing up to 8 MB of data. Data from the current flight overwrites older information when the memory section is full. The DSU stores the following information:
Engine Configuration Selector
•
Engine identification data
•
Trim balance data
•
Engine snapshot data
The engine configuration selector is used by EEC software to select control laws parameters compatible with the engine hardware configuration.
•
Engine historical data
DSU Failures
Engine configuration and identification data is programmed on the DSU during production acceptance testing and is not modifiable on the aircraft. The DSU remains with its specific engine for the life of the engine, unless the DSU fails.
If the DSU is not installed, and the aircraft is on the ground, the EEC does not allow the engine to start. If the DSU disconnects in flight, the EEC uses the programming information stored in the EEC memory to continue controlling the engine.
Engine identification data is read from the DSU during EEC initialization. Each channel performs an integrity check on the engine identification data in the DSU and in its own data flash. The DSU as the primary source, and uses its own data as a back-up source.
DSU critical faults such as nameplate and aircraft type incompatibility with the selectable engine ratings in one or both EEC channels result in an ENGINE FAULT advisory message and an 73 L(R) ENGINE FAULT FADEC FAULT 1 INFO message. The aircraft cannot be dispatched.
The engine serial number is used by the prognostics and health monitoring unit (PHMU) for engine condition monitoring.
A DSU noncritical fault such as a failure to write data to the DSU detected by one channel only results in an ENGINE FAULT advisory message and an 73 L(R) ENGINE FAULT - FADEC FAULT 2 INFO message.
N1 Modifier To ensure that all engines have uniform N1/thrust characteristics, an N1 class modifier is defined and its value is set in the DSU for each engine by tests during production. The EEC adds the N1 class modifier to the mechanical N1 speed to get the N1 feedback used in the N1 control loop and the N1 speed Indication displayed on EICAS.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The DSU dispatch failures are only annunciated on the ground to prevent an ENGINE FAULT advisory message and associated INFO message from being displayed in flight should a DSU failure be detected after an in flight shutdown.
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73 - Engine Fuel and Control 73-20 Control System
DATA STORAGE UNIT
DATA STORAGE UNIT DATA Engine Identification Data
8 MB Flash
Trim Balance Data
DSU Hardware Part Number
Fan Weight Data
DSU Data Part Number
Low Pressure Turbine Weight Data
Engine Serial Number Trim Balance Solution Data Engine Nameplate Thrust (PW1525G, PW1524G, PW1521G, PW1521GA, or PW1519G) 8 MB Flash
8 MB Flash
Channel A Processor
Channel B Processor
Engine Snapshot Data
Engine Historical Data
Engine Faults (such as sensor or actuator failures)
Total Engine Running Time
Engine Events (such as exceedances and surges)
Start Counter
Engine Run Time in Flight
Shutdown Counter
Vibration History Data Aborted Start Counter Data for Trim Balance Calculations
Number of Flights
Aircraft Type (CS100, or CS300)
Preferred Channel in Control
N1 Modifier
Preferred Igniter
CRC32 Checksum
Next Overspeed Shutdown Test Set-up (channel A, channel B, or channel A and B)
ELECTRONIC ENGINE CONTROL
Overspeed Shutdown Test Failed Journal Oil Shuttle Valve Shutdown Test Failed
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
NVM Engine Position
CS1_CS3_7321_001
HPT Disk Temperature at Shutdown
LEGEND CAN BUS
Figure 18: Data Storage Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
CONTROLS AND INDICATIONS THROTTLE QUADRANT ASSEMBLY The throttle quadrant has two thrust levers to control the engines. Each thrust lever controls the forward and reverse thrust functions. The L(R) ENG switches provide engine run and shutdown commands. ENGINE PANEL The ENGINE panel CONT IGN PBA provides a manual start request signal to the EEC for engine starting. The START switch is spring loaded to the AUTO position for normal automatic starts. The switch has L(R) ENG CRANK positions for manual engine starts. ENGINE FIRE PANEL The ENGINE FIRE panel FIRE PBA provides a shutdown signal to the EEC when pressed. The FIRE PBA is used to shut down the engine should the L(R) ENG switch shutdown command fail.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
ENGINE CONT IGNITION
MAX
MAX
IDLE
IDLE
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
L ENG
R ENG
ON
ON
FF FF OFF
FF FF OFF
THROTTLE QUADRANT ASSEMBLY
L ENG
R ENG
F I RE FIRE E
F IRE IR E
BTL 1
BTL 2
BTL 1
BTL 2
AVAIL
AVAIL
AVAIL
AVAIL
L ENGINE FIRE PANEL
R ENGINE FIRE PANEL
CS1_CS3_7320_005
ENGINE PANEL
Figure 19: Fuel Control System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
INDICATIONS The fuel control system indications are displayed on the engine indication and crew alerting system (EICAS) page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
CLB 92.2
92.2
73.0
73.3
N1
VIB
718
722
EGT
86.5 VIB N2 86.6 1090 FF (KPH) 1110 124 OIL TEMP 124 116 OIL PRESS 116 FAN VIB 0.2 0.4 CS1_CS3_7320_008
EICAS
Figure 20: Fuel Control System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
DETAILED DESCRIPTION AIRCRAFT AND ELECTRONIC ENGINE CONTROL INTERFACE
Primary Flight Control Computers
The following aircraft systems interface with the EEC for engine control.
The primary flight control computers supply a steep approach signal used to modify the engine idle speed for a steep approach.
Flight Management System The flight management system (FMS) provides flight plan information used to calculate the engine thrust reference setting. The FMS also supplies the thrust mode selection automatically and is the interface for manual selection of thrust mode, automatic power reserve (APR) arming, N1 synchronization arming, and FLEX temperature selection.
Ice Detection The EEC uses the ice detector signal to control the cowl anti-ice valve (CAIV) operation. Engine Panel The manual or automatic engine start mode is selected from the ENGINE panel.
Integrated Air System Controllers The integrated air system controllers (IASCs) supplies the EEC with environmental control system (ECS) and wing anti-ice system configuration. The IASCs receive bleed status information from the EECs. Air Data Information The EEC uses aircraft air data as primary source of air data because of its accuracy and redundancy. The air data system supplies total air temperature, pressure, and Mach information. The air data system receives T2 information for use in the event of a TAT failure.
Autothrottle The EEC provides the N1 target command to the autothrottle system. The autothrottle controls the thrust levers based on the N1 target information. The autothrottle also provides engine trim signals for engine synchronization. ENGINE FIRE Panel The ENGINE FIRE panel supplies engine shutdown commands when an engine fire PBA is pressed.
Landing Gear and Steering Control Units
Throttle Quadrant Assembly
The EEC use the weight-on-wheels (WOW) signal for engine idle speed control, engine control, and thrust reverser control.
The throttle quadrant assembly (TQA) provides the EEC with thrust lever position information and engine ON/OFF commands. The EEC supplies a signal to energize the baulk solenoid when the thrust reverser translating sleeves are deploying or stowing.
Brake Data Concentrator Units The brake data concentrator units (BDCUs) supply wheel speed signals for control of the thrust reversers.
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73 - Engine Fuel and Control 73-20 Control System
FLIGHT MANAGEMENT SYSTEM Thrust Reference Flex Temperature APR Arming N1 SYNC Thrust Mode Selection
THROTTLE QUADRANT ASSEMBLY
Start Mode Thrust Lever Angle Engine ON/OFF
IASC 1 and IASC 2
Baulk Solenoid Command
ENGINE PANEL
ECS Configuration Wing Anti-ice Configuration Bleed Status
Autotrim N1 Target
AIR DATA SYSTEM
AUTOTHROTTLE
Total Air Temperature Pressure Mach DMC 1 and DMC 2
LGSCU 1 and LGSCU 2
ELECTRONIC ENGINE CONTROL
Weight-On-Wheels
BDCU 1 and BDCU 2 Ice Detected ICE DETECTOR PFCCs Steep Approach
Engine Shutdown Command LEGEND
ENGINE FIRE PANEL BDCU DMC IASC LGSCU PFCC
Brake Data Concentrator Unit Data Concentrator Unit Module Cabinet Integrated Air System Controller Landing Gear and Steering Control Unit Primary Flight Control Computers ARINC 429
CS1_CS3_7320_002
Wheel Speed
Figure 21: Aircraft and Electronic Engine Control Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
ENGINE SENSOR INTERFACE
Pamb Sensor
The engine sensors are used to provide fuel control and engine health monitoring.
The single-channel Pamb sensor measures the ambient pressure through a port on EEC. The Pamb sensor is connected to channel A of the EEC. Channel B receives Pamb data via the EEC CAN BUS.
P2/T2 Probe The P2/T2 probe consists of a single channel P2 probe that connects to channel B of the EEC and a dual channel T2 sensor. The P2 probe measures the engine inlet total pressure. P2 is used to validate the air data from the air data system, control scheduling functions, engine rating calculation and Mach number calculation. The P2 probe is heated. The left P2/T2 probe heater is powered by AC BUS 1 and the right P2/T2 is powered by AC BUS 2. The EEC commands the probe heat on when the engine is running above idle and the temperature is below 9°C (48°F). When the probe heat is on, the EEC software compensates for the effect of the probe heater. The EEC turns the P2/T2 heater off when the engine is shutdown or the temperature is above 9°C. If the EEC detects a heater failure, a 73 L ENGINE FAULT - P2/T2 HEATER INOP INFO message is displayed. The T2 sensor measures the engine total inlet temperature. The T2 sensor feeds both channels of the EEC. T2 is used to validate the total air temperature information from the air data system, control scheduling functions including engine rating calculation and Mach number calculation. The T2 sensor information is used in place of air data information when the aircraft is on the ground and below 60 kt. The P2T2 probe is used if air data from the air data system is not valid or not available. If the P2T2 probe also fails, the EEC uses default values: •
Inlet total temperature: synthesized total temperature based on standard day temperature, Mach and pressure
•
Total pressure: 14.7 psi
•
Pressure altitude: 17000 ft
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The engine sensed altitude is calculated based on Pamb. Pamb is used to validate the air data from the air data system, control scheduling functions, engine rating calculation and Mach number calculation. P2.5/T2.5 Sensor The P2.5/T2.5 sensor consists of a pressure probe and a temperature sensor. The P2.5 pressure probe measures the pressure at the low-pressure compressor (LPC) exit. The probe receives an excitation signal from the PHMU and provide an output used to monitor high-pressure compressor (HPC) and LPC health. The T2.5 temperature sensor measures total temperature at the LPC exit. The sensor receives excitation from channel A of the EEC and provides an output used to monitor high-pressure compressor (HPC) and LPC health. Channel B of the EEC receives T2.5 sensor data via the EEC CAN BUS. T3 Sensor The T3 thermocouple sensor provides a single output to channel B of the EEC. It measures the HPC exit temperature. T3 is used to assess engine performance and compressor health. Pb Sensor The Pb (burner pressure) sensor is a dual channel pressure sensor. The EEC provides excitation voltage to the sensor. Each channel has a low and a high range pressure transducer. Pb is used in fuel scheduling, stall detection, autostart logic, continuous ignition, and shaft shear detection.
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73 - Engine Fuel and Control 73-20 Control System
P2.5/T2.5 PROBE P2.5 T2.5
ELECTRONIC ENGINE CONTROL T3 SENSOR PROGNOSTICS AND HEALTH MANAGEMENT UNIT
Channel A
Channel B Pb SENSOR Channel A
PAMB
P2 Channel B
P2/T2 PROBE T2 CDC 1-15-7 L P2/T2 HTR AC SSPC 3A BUS 1
P2
LEGEND Pb
Air Pressure Burner Pressure CAN BUS
NOTE Left engine shown. Right engine similar.
CS1_CS3_7320_006
Logic: Single Input
Figure 22: Engine Sensor Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
ENGINE OPERATING LIMITS The engine operating limits ensure that the engine operates within the design parameters. The EEC uses feedback from engine sensors and inputs from aircraft systems to ensure that the design limits are not exceeded. A governing loop sets the primary engine thrust requirements using N1 as a reference. The governing loop is set based on aircraft system inputs for the thrust requirements. The N1 is limited to maximum speed of 10,600 rpm (100%) and a minimum of 1,616 rpm (15%). The minimum N1 limit is dependant on environmental conditions, minimum thrust level, ice accumulation and bleed air requirements. The governing loop is constrained by EEC limiting loops.
The fuel flow to burner pressure ratio limit loop makes sure that the fuel to air ratio inside the combustor is maintaining a margin against flameout. The burner pressure maximum limit loop protects the engine from over pressurization of the combustion chamber. The burner pressure minimum limit loop also ensures enough pressure for the environmental control system. The minimum burner pressure depends on the demands for engine air made by the following systems: •
Cowl anti-ice
•
Wing anti-ice
•
Environment control system
The fuel flow rate maximum and minimum limit loops are provided to prevent excessive fuel command changes during acceleration and deceleration. During engine starting, the EEC modifies the fuel schedule once the starter air valve (SAV) closes and the air turbine starter (ATS) disengages.
The following limiting loops are monitored by the EEC: •
N2
•
Fuel flow
•
Fuel flow/burner pressure ratio
•
Burner pressure
Thrust Limitation at Low Airspeed
•
Fuel flow rate
A thrust limitation at low air speed (TLLS) function is available in TO mode only. TLLS maintains a minimum nose landing gear loading during the takeoff phase.
The N2 is limited to maximum speed of 24,470 rpm (100%) and a minimum of 13,087 rpm (53%). The minimum N2 limit ensures that the engine remains above generator trip speed. If N1 or N2 increase beyond 100%, but do not reach the overspeed shutdown condition of 105%, a L(R) ENG EXCEEDANCE caution message is displayed and the engine thrust must be reduced manually using the thrust levers. The fuel flow maximum limit protects the IFPC. There is a fuel flow minimum that assures a minimum amount of fuel is delivered to the combustor during starts. While in flight, the minimum fuel flow provides flameout protection at high altitude and during high G maneuvers. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The TLLS function is enabled when the aircraft is on the ground and the airspeed is less than 80 kt. As the airspeed increases, the TLLS input is reduced and the required thrust commanded by the EEC is provided by 80 kt. The TLLS function is inhibited when the aircraft is on the ground and the parking brake is set. This ensures full power can be developed during a high-power ground run.
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73 - Engine Fuel and Control 73-20 Control System
ENVIRONMENT CONTROL SYSTEM
DMC 1 and DMC 2
AIRCRAFT SYSTEMS
WING ANTI-ICE SYSTEM
EEC N1 SPEED SENSOR
PMAG (N2)
Governing Loop • N1 Limiting Loops • N2 • Fuel Flow • Fuel Flow/Burner Pressure Ratio • Burner Pressure • Fuel Flow Rate
BURNER PRESSURE SENSOR
COWL ANTI-ICE
STARTER AIR VALVE
AIR TURBINE STARTER
IFPC DMC EEC IFPC PMAG
Data Concentrator Unit Module Cabinet Electronic Engine Control Integrated Fuel Pump and Control Permanent Magnet Alternator Generator ARINC 429
Fuel Metering Valve NOTE LH Engine shown, RH similar.
CS1_CS3_7321_033
LEGEND
Figure 23: Electronic Engine Control Engine Operating Limits Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
OVERSPEED PROTECTION, SHAFT SHEAR PROTECTION AND THRUST CONTROL MALFUNCTION Overspeed protection
Thrust Control Malfunction
Overspeed protection is provided for both high (N2) and low (N1) pressure rotors. This is done via dedicated protection computers within the EEC independent of the channel A and B control computers.
A thrust control malfunction is an uncontrollable high thrust event. A thrust control malfunction is characterized by:
The protection computers control the overspeed shutdown solenoid (OSS). When either N1 or N2 engine speeds exceed the overspeed limit, at any time the engine is running, the engine immediately shuts down.
•
Uncommanded engine thrust increase
•
Engine thrust remaining at a high thrust setting when a lower thrust setting is commanded
The protection computers share the N1 and N2 signals provided to the control computers within the EEC and have a separate crosstalk CAN BUS.
In either case, the thrust can not be controlled by the thrust lever. When the EEC detects a thrust control malfunction, the OSS is energized to shut down the engine
The EEC limits fuel flow when redline limits are reached for N1 or N2. When either of the protection computer detects an overspeed limit of 105% N1 or 105% N2, the OSS is energized.
Thrust control malfunction detection is inhibited during:
Once an overspeed limit exceedance is detected, the system is latched and must be reset. The system can be reset by cycling the power supply to the EEC, or by selecting the L(R) ENG switch to ON to initiate the engine start sequence. To ensure that there are no undetected faults, the OSS is tested at shutdown.
•
Thrust reverse operation
•
Approach
•
Steep approach
Shaft Shear Protection The protection computers also provide protection against low and high pressure rotor shaft shear. If the protection computers determine that either shaft shear conditions have occurred, the OSS is energized to shutdown the engine.
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73 - Engine Fuel and Control 73-20 Control System
L ENG
R ENG
ON
ON
OFF
OFF
THROTTLE QUADRANT ASSEMBLY
INTEGRATED FUEL PUMP AND CONTROL
ELECTRONIC ENGINE CONTROL Channel A
N1 SPEED SENSOR
Control Computer
PERMANENT MAGNET ALTERNATOR GENERATOR
Overspeed Shutdown Solenoid
Channel B
Protection Computer
CS1_CS3_7321_034
Control Computer
N2
Protection Computer
LEGEND CAN BUS
Figure 24: Overspeed, Shaft Shear Protection, and Thrust Control Malfunction Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL OPERATION - START The IFPC is controlled by the EEC to provide metered fuel for combustion based on thrust lever position and ambient operating conditions. The IFPC also provides servo fuel for the fuel actuated system on the engine. The IFPC consists of both hydromechanical and electromechanical valves to control the fuel flow. The fuel metering valve (FMV) is a dual channel component controlled by the EEC. A dual channel torque motor allows either channel of the EEC to control the FMV. FMV feedback is provided to the EEC using a dual coil LVDT. The pressure-regulating valve (PRV) maintains a constant pressure drop across the FMV by returning excess pump flow to the inlet of the high-pressure pump. The excess flow is dependent on the servo fuel demand and metered fuel requirements. The fuel flow meter provides the EEC with an indication of the amount of fuel being used for combustion. When the engine is commanded to shutdown, the windmill bypass valve (WBV) opens to connect fuel from upstream of the minimum pressureshutoff valve (MPSOV) with the fuel filter inlet. The minimum pressure-shutoff valve (MPSOV) is a spring-loaded valve that prevents fuel from entering the primary and secondary fuel manifolds until it is sufficiently pressurized. The overspeed shutdown solenoid (OSS) is a dual-coil solenoid valve that operates the WBV and MPSOV to shutoff metered flow to the engine. The flow divider valve (FDV) schedules the fuel flow to primary and secondary manifolds depending on engine start requirements. The FDV is normally spring-loaded to the closed position, blocking fuel to the secondary fuel manifold. Once the fuel pressure between the primary and secondary fuel manifolds exceeds 100 psi, the FDV opens allowing fuel into the secondary manifold.
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73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL From Servo Pressure Regulator MAX
MAX
IDLE
IDLE
To Fuel Filter
OSS
Pressure-Regulating Valve
SOL
WBV L ENG
R ENG
ON
ON
OFF
OFF
FMV
TQA
Channel A
From Main Pump
EEC
Channel B Channel A
T/M SOL
EBV SOLENOID Channel B MPSOV
SOL T/M
EBV
FUEL/OIL MANIFOLD
To Primary Fuel Manifold To Secondary Fuel Manifold FUEL FLOW METER CS1_CS3_7321_039
EBV EEC FDV FMV MPSOV OSS TQA WBV
Metered Fuel Pressurized Fuel Return Fuel Servo Fuel Equalization Bypass Valve Electronic Engine Control Flow Divider Valve Fuel Metering Valve Minimum Pressure-Shutoff Valve Overspeed Shutdown Solenoid Throttle Quadrant Assembly Windmill Bypass Valve CAN BUS Solenoid Torque Motor
FDV
Figure 25: Integrated Fuel Pump and Control Operation - Engine Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL OPERATION - RUN The EEC energizes the EBV solenoid when the engine start is completed. The equalization bypass valve solenoid controls the equalization bypass valve (EBV). The EBV energizes to equalize the fuel flow in the primary and secondary fuel manifolds. The FDV closes and the secondary fuel manifold is supplied through the EBV.
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73-52
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL From Servo Pressure Regulator MAX
MAX
IDLE
IDLE
To Fuel Filter
OSS
Pressure-Regulating Valve
SOL
WBV L ENG
R ENG
ON
ON
OFF
OFF
FMV
TQA
Channel A
From Main Pump
EEC
Channel B Channel A
T/M SOL
EBV SOLENOID Channel B MPSOV
LEGEND
SOL T/M
EBV
FUEL/OIL MANIFOLD
To Primary Fuel Manifold To Secondary Fuel Manifold FUEL FLOW METER CS1_CS3_7321_008
EBV EEC FDV FMV MPSOV OSS TQA WBV
Metered Fuel Pressurized Fuel Return Fuel Servo Fuel Equalization Bypass Valve Electronic Engine Control Flow Divider Valve Fuel Metering Valve Minimum Pressure-Shutoff Valve Overspeed Shutdown Solenoid Throttle Quadrant Assembly Windmill Bypass Valve CAN BUS Solenoid Torque Motor
FDV
Figure 26: Integrated Fuel Pump and Control Operation - Engine Run Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-53
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
NORMAL ENGINE SHUTDOWN On Ground
In Flight
The EEC commands shutdown of the engine using the overspeed shutdown solenoid (OSS) whenever one of the following conditions is present:
In flight the EEC commands the engine to shutdown by driving the FMV to the minimum flow position first to allow time for the high-pressure compressor (HPC) stator vane actuator to track to the optimal start position.
•
L(R) ENG SWITCH is moved to the OFF position on the throttle quadrant assembly
•
An excessive engine overspeed is detected by the EEC
•
A shaft shear is detected by the EEC
When the HPC stator vane actuator is in position, the fuel metering valve is commanded closed and the OSS is energized.
The EEC uses the overspeed protection system to shut down the engine. The EEC channel in control commands the overspeed shutoff solenoid to energize. On every third shutdown both EEC channels command the overspeed shutoff solenoid to fully test the system. When the EEC energizes the OSS, servo fuel pressure opens the WBV to allow fuel to flow back to the pressure-regulating valve. The pressureregulating valve opens so that the fuel flows back to the fuel filter and is recirculated. At the same time, the MPSOV moves to the closed position and stops fuel from flowing to the engine fuel nozzles. The EEC de-energizes the equalization bypass valve solenoid to the closed position. Once the overspeed protection system has shut down the engine, or if the overspeed protection system fails to shut down the engine, the EEC commands the fuel metering valve closed. When the fuel metering valve is confirmed closed, the overspeed shutdown solenoid is de-energized and all EEC processors are reset.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-54
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
INTEGRATED FUEL PUMP AND CONTROL From Servo Pressure Regulator MAX
MAX
IDLE
IDLE
To Fuel Filter
OSS
Pressure-Regulating Valve
SOL
WBV L ENG
R ENG
ON
ON
FF FF OFF
FF FF OFF
FMV
TQA
Channel A
From Main Pump
EEC
Channel B Channel A
T/M SOL
EBV SOLENOID Channel B MPSOV
SOL T/M
FUEL/OIL MANIFOLD
EBV
To Primary Fuel Manifold To Secondary Fuel Manifold FUEL FLOW METER CS1_CS3_7321_038
EBV EEC FDV FMV MPSOV OSS TQA WBV
Metered Fuel Pressurized Fuel Return Fuel Servo Fuel Equalization Bypass Valve Electronic Engine Control Flow Divider Valve Fuel Metering Valve Minimum Pressure-Shutoff Valve Overspeed Shutdown Solenoid Throttle Quadrant Assembly Windmill Bypass Valve CAN BUS Solenoid Torque Motor
FDV
Figure 27: Normal Engine Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-55
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
EMERGENCY SHUTDOWN An emergency shutdown can be carried out using the L ENG or R ENG FIRE PBA. Pressing the fire PBA initiates the following: •
Engine feed shutoff valve closes
•
Pressure-regulating shutoff valve (PRSOV) closes
•
Fan air valve closes
•
Hydraulic shutoff valve closes
•
Generator trips offline
•
EEC commands engine shutdown
•
Both engine fire extinguisher bottles are armed
The integrated air system controller (IASC) and electronic engine control (EEC) receive digital inputs from the FIRE PBA via an ARINC 429 BUS. The fan air valve and pressure-regulating shutoff valves are closed by the IASC. The EEC provides a shutdown command to the engine. The engine feed shutoff valve is closed through switch contacts in the FIRE PBA.
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73-56
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
INTEGRATED COCKPIT CONTROL PANEL Left Engine Fire Panel BTL 1
L ENG
AVAIL FIRE
EPC 1 BATT DIR BUS 1
5
Firex A
To Fire Extinguisher 1
EPC 2 BATT DIR BUS 2
5
BTL 2
Firex B
K1
AVAIL
To Fire Extinguisher 2 From L GEN PBA
K2
L Generator Offline
NOTE
PIM
Left engine shown. Right engine similar. LEGEND Data Concentrator Unit Module Cabinet Electronic Engine Control Electrical Power Center Integrated Air System Controller Panel Interface Module ARINC 429
L IASC
L EEC
LEFT ENGINE FEED SHUTOFF Closed VALVE Open
HYD SOV Closed
CS1_CS3_7321_031
DMC EEC EPC IASC PIM
DMC 1 and DMC 2
Figure 28: Emergency Shutdown Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-57
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
ENGINE POWER MANAGEMENT Engine Bleed Management Each EEC receives information from the integrated air system controllers (IASCs). IASC 1 provides information to the left EEC. IASC 2 provides information to the right EEC. •
High-pressure valve (HPV) open or closed status
•
Environmental control system (ECS) configuration
•
Anti-ice configuration: -
Wing anti-ice status
-
Nacelle anti-ice status
Idle Speed The EEC selects engine idle speed based on engine requirements and bleed demand. The minimum idle speed is based on the following priorities: •
Minimum fuel flow for combustor stability
•
Minimum pressure to provide the required aircraft bleed flow, and maintaining the minimum pressure to not exceed a 20% core flow bleed limit
•
Minimum N1 to achieve acceleration time requirements
•
Minimum N2 speed to maintain the PMAG power generation
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
IASC INPUTS
High-Pressure Bleed Port • Closed • Open
EEC IDLE SETTING DECISION TREE
1. Minimum FUEL flow for combustor stability.
2. Minimum pressure to provide aircraft bleed flow (not to exceed 20% of the core flow).
Select Highest Value
Set Idle
ECS Bleeds • No Flow • Low-Flow • High-Flow 3. N1 Idle for transient acceleration times.
Wing Anti-Ice • No Flow • Low-Flow • High-Flow
LEGEND Environmental Control System Electronic Engine Control Integrated Air System Controller
CS1_CS3_7321_029
ECS EEC IASC
4. N2 Idle to maintain minimum generator speed.
Figure 29: Engine Bleed Management and Idle Speed Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-59
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
ENGINE POWER SETTINGS AND MODES Using flight plan performance data from the flight management system (FMS), environmental data, and engine sensors, the EEC calculates the N1 target for maximum takeoff and go-around, maximum continuous thrust, derated takeoff, reduced takeoff (FLEX), and climb thrust. Maximum Thrust The EEC sets the maximum thrust available at the MAX thrust position to maximum takeoff thrust (MTO) when the aircraft is operating in the takeoff or go around envelope. Outside the takeoff envelope or go around envelope, the EEC schedules maximum continuous thrust (MCT). The MCT rating supports one engine inoperative (OEI) conditions and is designed to be lower or equal to MTO thrust and higher than or equal to climb thrust (CLB). Mode Selection The thrust lever is adjusted automatically to the selected thrust mode rating when the autothrottle is engaged. Thrust mode changes are managed automatically based on the FMS active flight plan. The thrust mode can be controlled manually on the FMS PERF page DEP tab using the THRUST button. A window pops up to enable manual selection. Derated Takeoff Derated takeoffs TO-1, TO-2, or TO-3 are used to reduce fuel consumption, noise, and to extend engine life and reliability. The derated TO thrust selection is computed using the mode selection on the FMS DEP page. The takeoff derate is limited to a minimum of 17,000 pounds of thrust.
the assumed temperature input is higher than the actual temperature, the FLEX reduction in thrust is applied for that takeoff. FLEX is available from any selected takeoff thrust mode. Climb Thrust The FMS calculates the optimal climb (CLB) thrust and N1 reference based on the takeoff thrust mode (derated, FLEX), and the outside air temperature (OAT). Derated climb CLB-1 or CLB-2 allow the thrust settings for the climb envelope of the flight to be reduced by the FMS. There are two levels of climb derate. Each subsequent level reduces the thrust during the climb. Go-Around Go-around thrust is set as a thrust reference when the aircraft is on approach with the slats and flaps extended and the landing gear down. Go-around thrust is equivalent to MTO. Go-around thrust is initiated when the aircraft is in flight and the takeoff/go-around (TOGA) switches on the thrust levers are pressed. Idle Speed The EEC determines the idle setting based on engine stability requirements, bleed demand, and phase of flight. The minimum idle speed setting minimizes the aircraft landing distance and brake wear during taxi. Flight idle is used in flight to minimize fuel consumption during descent. When the landing gear is selected down or the flaps are extended, the engine idle speed increases to the approach idle setting to allow a faster engine response during a go-around. During a steep approach, a lower than normal approach idle setting is used. The steep approach idle is adjusted for bleed air demands and goaround thrust requirements.
FLEX Thrust FLEX thrust is computed using the assumed temperature method, based on the value entered in the FMS. On the ground, when the EEC confirms Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-60
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
Departure Tab
ACT
FMS1
POS
CLB
DEP Flex
DBASEE DBAS
FPLN
CR CRZ RZ
DES
ORIGIN
RWY
SID
LFBO
RW06L
CHATY2 ------
TO THRUST
FLEX
OAT
TO-1
+15
°C
+10
PERF
ROUTE ROUT UTE ARRR
RWY WIND °C
TO-3
Performance Page Selection
TRANS
FLEX 44°C
CLB THRUST
080°// 14
73.3
73.3
73.3
73.3
CLB-1 N1
APR
ARMED Fixed Takeoff Derate Selection
V1
EICAS
Fixed Climb Derate Selection
V2
SET VSPEEDS DS
123 125 130
APR Selection
COMPLETE F3
3
150
BLEEDTHRUST BLEE SOURCE
APU AP
F2 160
F1 170
180
WING ANTI-ICE
MODE TRANS
CLB THR CLB-1 CLB-2
Symbol
GD 210
Manually entered thrust reference value.
OFF OPT ALT MAX ALT STAB TRIM FL330 FL410 ND 6.2
PITCH SPD THRUST OEI FMS PERF VFLC FMS CLB TO
80 %N1 VFLC RED FLAPS CLEAN %N1
Description FMS managed thrust reference value.
COWL ANTI-ICE
OFF
CRZ ALT AUTO TOW TOW CG FL250 5500 LB AUTO 55000 33.5 %MAC TO-1 + APR PR 100 %N1 DEP PROFILE MAN HT/ALT 100 %N1 GA / 1618 STANDARD MCT ACCEL 1500 94 ST %N1 Thrust Button
F0
CNCL C NC L NCL
FMS CLB CLB
ENGINE SYNC
%N1
DONE D DO O NE
THRUST...
Thrust mode. TO-3 Assumed temperature for FLEX XXºC. FLEX 44°C 73.3 73.3 Digital thrust reference.
TO TO-1 TO-2 TO-3 CLB CLB-1 CLB-2 MCT GA
or or or or or or or or or
TO TO-1 TO-2 TO-3 CLB CLB-1 Magenta: FMS managed CLB-2 Cyan: manual selection MCT GA
FLEX 44°C 73.3 73.3 FLEX 44°C 73.3 73.3
FMS PERFORMANCE PAGE
CS1_CS3_7321_036
SLAT/FLAP Auto Manual Selection
VR
Figure 30: Engine Power Settings and Modes Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-61
CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
AUTOMATIC POWER RESERVE
Scenario 3
The automatic power reserve (APR) automatically provides a higher thrust setting to the operating engine should one of the engines fail during a derated takeoff. APR provides improved single-engine acceleration and climb gradient performance.
The engines are derated to TO-3 and FLEX is selected. When the EEC detects one engine has lost thrust, the derate changes to TO-2 and FLEX is canceled. The operative engine is commanded to supply TO-2 thrust.
APR is only available when a derated takeoff thrust reference is selected. APR is not available for any other thrust mode. Should one engine fail, the EEC automatically increases the thrust on the operative engine by canceling the engine fixed derate by one level and removing any FLEX entry. When the thrust increase is commanded, the thrust levers angle (TLA) does not change.
Additional thrust is available by advancing the thrust levers. The EEC recalculates the TLA versus N1 command relationship to ensure the full range of thrust is available from the engine.
If TO-1 derate is selected, the engine thrust increases to MTO. If TO-2 or TO-3 are selected, the thrust is commanded to a level below MTO in order to maintain aircraft controllability. MTO is available by advancing the thrust levers to MAX. When armed, APR activates when both thrust levers are advanced beyond 23° and one engine experiences a loss of power greater than 15% N1. The APR can be manually disarmed from the FMS performance page DEP tab. When APR is manually disarmed selected, an APR DISARM status message is displayed. Scenario 1 The engines are derated to TO-1 prior to takeoff. This cancels one fixed derate level and maximum thrust is commanded on the operative engine. Scenario 2 The engines are derated to TO-1 and FLEX is selected. When the EEC detects one engine has lost thrust, the derate and FLEX are canceled. Maximum thrust is commanded on the operative engine.
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TECHNICAL TRAINING MANUAL
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
LEGEND APR MTO TLA TO
Thrust
Automatic Power Reserve Maximum Takeoff Thrust Lever Angle Takeoff
B2
SCENARIOS
A2
23,300 lbf
Scenario 1
A1 to A2
The derate is canceled and thrust is increased to MTO.
A1
21,000 lbf
Scenario 2
B1 to B2
Both the derate and the FLEX are canceled, thrust is increased to MTO.
C1 to C2
The derate is reduced to TO-2 and FLEX are canceled. Engine thrust automatically increases to TO-2. Additional thrust is available by advancing the thrust levers.
TO-1
18,900 lbf
C2
Scenario 3
17,000 lbf APR Activated
B1 C1
TO-2
TO-1 + FLEX
TO-3 + FLEX
88.5
APR 88.5
20.0
88.5
N1 EICAS PAGE
TO THRUST
0° TO-3
TO-2
TO-1
TO (47.5°)
TO-1
+15
°C
APR
ARMED
THRUST LEVER ANGLE VERSUS APR RELATIONSHIP
FMS PERFORMANCE PAGE
CS1_CS3_7321_030
Thrust Lever Angle
FLEX
Figure 31: Automatic Power Reserve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
N1 SYNCHRONIZATION The N1 synchronization (N1 SYNC) function is used to synchronize the engine speeds to reduce cabin noise. The N1 SYNC function is selectable through the aircraft flight management system (FMS) and is armed by default. The N1 SYNC function maintains the difference between the sensed N1 values of both engines to less than 0.2% during steady state conditions. N1 SYNC is enabled by the EEC, provided specific engine operating conditions are satisfied. The left engine is the master engine for the N1 SYNC function. The N1 command for the slave engine is biased to match that of the master engine. With the N1 SYNC mode set in the FMS, the EEC activates N1 SYNC if the following conditions are met: •
Both engines are operating in a steady-state
•
The aircraft is not in takeoff, or go-around
•
The difference between the two engines N1 commands is less than 1.5% N1
•
Throttle lever angle (TLA) is greater than 6°
•
No opposite engine low-thrust condition or automatic power reserve selection is detected
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73 - Engine Fuel and Control 73-20 Control System
THRUST MODE TRANS CRZ ALT FL250
AUTO AUTO TO-1 + APR MAN GA
MCT CLB CLB-1 CLB-2
OPT ALT FL330
MAX ALT FL410
100 %N1 100 %N1 94 %N1
OEI FMS PERF
CCNCL NCL NCL
80 %N1 %N1
ENGINE SYNC
%N1
DONE D DO ONE ONE FMS PERFORMANCE PAGE
DATA CONCENTRATOR UNIT MODULE CABINET
N1 SYNC Command
Electronic Engine Control
92.2
92.2
73.0
73.2
N1 SYNC EICAS
Integrated Fuel Pump And Control LEFT ENGINE
Integrated Fuel Pump And Control
LEGEND ARINC 429
RIGHT ENGINE
CS1_CS3_7321_028
Electronic Engine Control
CLB FLEX 44°C
Figure 32: N1 Synchronization Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_73_L3_ControlSystem_DetailedDescription.fm
73 - Engine Fuel and Control 73-20 Control System
FAULT REPORTING AND TESTING INTERFACE
-
No dispatch
Each EEC channel has a non-volatile memory (NVM) that can store snapshots and transient records of parameters. Each EEC channel records its own faults or events that are detected. Old data is overwritten by new data when the memory is full.
-
Short time dispatch
-
Long time dispatch
The PHMU transmits data to the health management unit (HMU) via the EEC. The EEC sends the data via an ARINC 429 bus for distribution to the health management unit (HMU) and the onboard maintenance system (OMS). The EEC also provides information to generate CAS and INFO messages on EICAS and the INFO synoptic pages. The CAS and INFO messages are used to determine the capability to dispatch the aircraft. The health management unit is capable of recording the following: •
Fault data such as failures detected within FADEC system components or its interfaces with the aircraft such as sensor and actuator failures
•
Event data consisting of abnormal operation such as surges and exhaust gas temperature exceedances
•
Health trending data consisting of parameters collected on every flight to track engine performance over time. The data is used for trend monitoring and can provide early detection of an engine failure. The data can also be used to generate reports such as takeoff reports and stable cruise reports
The OMS records faults annunciated by the EEC. The OMS also provides the capability of testing EEC functions and displaying real-time data that support line maintenance operations. (see ATA 45). Dispatch The EEC monitors all faults and classifies them according to the operability of the engine. Faults are classified as follows: •
•
Master minimum equipment list (MMEL) faults
Time limited dispatch faults include: •
Failures directly used for thrust control
•
Malfunctions resulting in an auto-idle or auto-shutdown condition
•
Malfunctions resulting in a thrust lever moved to idle or an engine shutdown in response to a CAS message
No dispatch faults are major failures that prevent operation or control of the engine. A no dispatch fault has an L(R) ENGINE FAULT advisory message and a 73 L (R) ENGINE FAULT - FADEC FAULT 1 INFO message. Short time dispatch faults usually relate to a single channel failure that can be compensated for by alternate means on the engine. Short time dispatch faults are limited to 125 flight hours. A short time fault has an L(R) ENGINE FAULT advisory message and a 73 L (R) ENGINE FAULT - FADEC FAULT 2 INFO message. Long time dispatch faults usually relate to a failure that does not affect the operability of the engine such as a sensor dedicated to engine performance monitoring. Long time dispatch faults are limited to 425 flight hours. A long time fault has an L(R) ENGINE FAULT advisory message and a 73 L (R) ENGINE FAULT - FADEC FAULT 3 INFO message. MMEL faults are normally related to secondary systems or indication system not essential for engine operation. MMEL faults are generally associated with caution messages.
Time limited dispatch
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73 - Engine Fuel and Control 73-20 Control System
Maintenance Main Menu
DMC 1 and DMC 2
View Post Flight Summary View Flight Deck Effects ( 28 Active) View Fault Messages ( 43 Active) View Service Messages ( 13 Active) View System Exceedances ( 0
Stored)
View System Trends
HEALTH MANAGEMENT UNIT
View Aircraft Life Cycle Data View System Parameters Perform LRU/System Operations
ELECTRONIC ENGINE CONTROL
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
R ENG OPER DEGRADED R ENGINE FAULT XXXXXXXXXXXXXXXXXXX XXXXXXXXXXXXXXXXXXX
View System Configuration Data Write Maintenance Reports
Utility Functions
OMS
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
73
R ENGINE FAULT - FADEC FAULT 2
30
L ENG ANTI ICE - LOW PRESS SW INOP
LEGEND DMC OMS
Data Concentrator Unit Module Cabinet Onboard Maintenance System ARINC 429 CAN BUS
PREVIOUSLY ACKNOWLEDGED 27
FLIGHT CONTROL FAULT - AILERON FORCE MON INOP
SYNOPTIC PAGE – INFO
CS1_CS3_7320_007
EICAS
Figure 33: Electronic Engine Control Fault Reporting and Testing Interface Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73 - Engine Fuel and Control 73-20 Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine controls system.
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CS130_73_L2_ControlSystem_MonitoringTests.fm
73 - Engine Fuel and Control 73-20 Control System
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L (R) ENG EXCEEDANCE Left (R) N1 or left (R) N2 or left ITT or left oil temperature above threshold. L (R) ENG OPER DEGRADED
Uncertainty on T2/TAT (degraded or default value), or HPC stator vane actuator not tracking as it should.
L (R) ENG FAIL
L 9R) engine Sub-idle.
ENG SETTINGMISMATCH
Thrust management data received or fed back by FADEC not matching the APR setting, N1 target, etc.
Table 2: ADVISORY Messages MESSAGE L (R) ENGINE FAULT
LOGIC Loss of redundant or non-critical function for the left engine.
Table 3: STATUS Messages MESSAGE APR DISARMED
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LOGIC APR is disarmed in the FMS and a derated takeoff is being used.
TECHNICAL TRAINING MANUAL
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73 - Engine Fuel and Control 73-20 Control System
Page Intentionally Left Blank
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73 - Engine Fuel and Control 73-20 Control System
Table 4: INFO Messages Table 4: INFO Messages MESSAGE
MESSAGE LOGIC
73 L (R) ENGINE FAULT - EEC A CTRL CPU INOP
EEC channel A control processor has failed.
73 L (R) ENGINE FAULT - EEC B CTRL CPU INOP
EEC channel B control processor has failed.
73 L (R) ENGINE FAULT - PT/T2 HEATER INOP
FADEC logic has determined that the P2/T2 heater has a logical failure and is indicated as failed.
73 L (R) ENGINE FAULT - FADEC FAULT 1
Collector bit for all non-dispatchable FADEC faults.
73 L (R) ENGINE FAULT - FADEC FAULT 2
Collector bit for all FADEC short term TLD faults.
73 L (R) ENGINE FAULT - FADEC FAULT 3
Collector bit for all FADEC long term TLD faults.
77 L (R) ENGINE FAULT - PHMU INOP
FADEC logic has determined that the PHMU is failed.
73 L (R) ENGINE FAULT - EEC 28VDC REDUND LOSS
Set when only one EEC channel of the left (right) engine has detected loss of 28 VDC input voltage.
73 L (R) ENGINE FAULT - HEALTH MON DEGRADED
Set by the left (right) engine when the P2.5 or T2.5 sensor is failed.
73 L (R) ENGINE FAULT - T3 SNSR INOP
Set by the left (right) engine when the T3 sensor is failed. This INFO message will be masked if the EEC Channel B control processor is failed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
73 L (R) ENGINE FAULT - FUEL EQUAL INOP
TECHNICAL TRAINING MANUAL
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LOGIC Set by the left (right) engine in any of the following conditions: Both EEC channels are unable to command fuel equalization solenoid. Any channel control processor failed and remaining channel unable to command fuel equalization solenoid.
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73 - Engine Fuel and Control 73-20 Control System
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM
System Interactive Test
The aircraft onboard maintenance system (OMS) uses the electronic engine control (EEC) to perform functional checks of specific systems. Once the new EEC is installed, the EEC SYSTEM TEST must be carried out. This test is performed through the OMS using one channel at a time. Each system test should be checked through both channels.
The system interactive test procedure is used to check for satisfactory operation of the EEC system before an engine start.
The following EEC tests can be performed through the aircraft OMS:
This procedure is used to verify the presence of existing faults, and to confirm that corrective action has eliminated the fault. The EEC system interactive test electrically verifies the EEC, and prognostics and health management unit (PHMU) input and output circuits.
•
Actuator Test
•
EEC System Test
•
HPD Test
•
Harness Test
•
Ignition Test
•
P2 Probe Heat Test
Harness Interactive Test
•
PCE Door Test
•
Thrust Reverser Cycling Test
The harness interactive test checks for harness and connector failures. During the test suspected harnesses must be moved to help reproduce the fault messages reported by the EEC.
NOTE The system interactive test does not test igniters, thrust reverser, or hydraulic pumps. Separate interactive tests are provided for these systems.
This procedure is used to perform harness checks in order to isolate intermittent EEC monitored faults. Probe Heat Test The P2/T2 probe heat test verifies the operation of the probe heater circuit. The EEC turns the probe heat on, and after 60 seconds verifies a rise in T2 temperature. The EEC then turns the probe heat off.
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73-72
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-20 Control System
MAINT MENU LRU/System Operations View
All ATAs
Select ATA
73 - ENGINE FUEL & CONTROL
Select LRU/System Left Engine EEC (1) channel A
DATA
RIGGING
TEST
NVM
Actuator Test EEC System Test HPD Test Harness Test Ignition Test P2 Probe Heater Test PCE Door Test
CS1_CS3_7330_003
Thrust Reverser Cycling Test
Figure 34: OMS EEC System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
73-73
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-20 Control System
HMU Power Assurance Test The power assurance test can be performed using the ATA 46 HMU CH A or CH B selection on the OMS LRU/System Operations page. Using ambient pressure, temperature, N1 errors, and nacelle anti-ice (NAI) status, the HMU software calculates and corrects the N2, EGT, and fuel flow parameter range. During the engine run, the HMU compares the actual engine parameters against the corrected parameter ranges. If all of the actual parameters are all within the corrected parameter ranges, a TEST PASSED message is displayed. A TEST FAIL message is shown if any parameter is out of range.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-74
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-20 Control System
MAINT MENU
MAINT MENU EMU-Engine 1 Power Assurance Test
LRU/System Operations View
All ATAs
Select ATA
46 - INFORMATION SYSTEMS 1/6 *** WARNING *** BOTH ENGINES WILL OPERATE DURING THE TEST. ENSURE TO FOLLOW THE ENGINE RUN-UP SAFETY PRECAUTIONS
Select LRU/System HMU CH A (A604)
DATA
TEST
NVM
*NOTE* TEST DURATION ~ 20 MINUTES
Engine 1 Power Assurance Test HMU - Cellular Connectivity Test HMU - IMS Interface Test
PRECONDITIONS
HMU - NVM Erase Function
1- REFER TO AMP XX-XX-XX-XX-XXXX 2- MAKE SURE THAT WIND SPEED < 35 KNOTS 3- MAKE SURE THAT PARKINBG BRAKE IS APPLIED 4- MAKE SURE THAT THE AIRCRAFT IS PROPERLY CHOCKED
HMU - WiFi Connectivity Function
TEST INHIBIT/INTERLOCK CONDITIONS OK
PRESS "CONTINUE" BUTTON TO GO TO NEXT PAGE PRESS "CANCEL" BUTTON TO GO TO MAIN MENU
Continue
Cancel
CS1_CS3_7330_004
PARKING BRAKE
Figure 35: Power Assurance Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-75
CS130_73_L2_ControlSystem_PracticalAspects.fm
73 - Engine Fuel and Control 73-30 Indicating
73-30 INDICATING GENERAL DESCRIPTION The fuel indicating system consists of a fuel flow meter to monitor fuel flow and a fuel filter differential pressure sensor to provide an indication of a blocked fuel filter. The fuel flow transmitter measures metered fuel for combustion. The EEC calculates the fuel flow for display on the engine indication and crew alerting system (EICAS) page. The fuel filter differential pressure sensor measures differential pressure across the fuel filter. The fuel filter status is displayed on the FUEL synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-76
CS130_73_L2_Indicating_General.fm
73 - Engine Fuel and Control 73-30 Indicating
CLB 92.2
92.2
73.0
73.3
N1
718
722 STATUS
EGT
FUEL FUE EL
86.5 1090 124 116
FUEL FLOW METER
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
TOTAL FUEL FUEL USED
AIR IR
DOOR OR
ELECC
FLT CTR CTRLL
HYD YD
AVIONIC IONIC
INFO
CB
15300 KG 1000 KG
EICAS
FUEL FILTER DIFFERENTIAL PRESSURE SENSOR
20 °C 2900 KG
20 °C 2900 KG
ELECTRONIC ENGINE CONTROL
9500 KG
APU
ARINC 429 SYNOPTIC PAGE – FUEL
CS1_CS3_7330_002
LEGEND
Figure 36: Fuel Control Indicating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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73-77
CS130_73_L2_Indicating_General.fm
73 - Engine Fuel and Control 73-30 Indicating
COMPONENT LOCATION The engine fuel indicating system consists of the following components: •
Fuel flow meter
•
Fuel filter differential pressure sensor
Fuel Flow Meter The fuel flow meter is located above the integrated fuel pump and control (IFPC) on the fuel/oil manifold at the 9 o’clock position. Fuel Filter Differential Pressure Sensor The fuel filter differential pressure sensor is mounted on the forward side of the fuel/oil manifold adjacent to the main fuel filter at the 7:30 o’clock position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-78
CS130_73_L2_Indicating_ComponentLocation.fm
73 - Engine Fuel and Control 73-30 Indicating
A
A
FUEL FLOW METER
B
FUEL FILTER DIFFERENTIAL PRESSURE SENSOR
CS1_CS3_7321_011
B
Figure 37: Fuel Control Indicating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-79
CS130_73_L2_Indicating_ComponentLocation.fm
73 - Engine Fuel and Control 73-30 Indicating
CONTROLS AND INDICATIONS The fuel flow information is displayed on the EICAS page. The fuel filter bypass status is indicated on the FUEL synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-80
CS130_73_L2_Indicating_ControlsIndications.fm
73 - Engine Fuel and Control 73-30 Indicating
CLB
STATUS
92.2
92.2
73.0
73.3
FUEL TOTAL FUEL FUEL USED
AIR
DOOR
ELEC
FLT CTRL
HYD
AVIONIC
INFO
CB
FUEL STATUS
9500 KG 3000 KG
Symbol
Condition Normal
N1
718
722 Fuel Filter Impending Bypass
EGT
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
5000 KG
Fuel Filter Bypass APU
BYPASS
SYNOPTIC PAGE – FUEL
CS1_CS3_7322_003
EICAS
23 °C 2000 KG
23 °C 2500 KG
Figure 38: Fuel Control Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-81
CS130_73_L2_Indicating_ControlsIndications.fm
73 - Engine Fuel and Control 73-30 Indicating
DETAILED DESCRIPTION The fuel flow meter sends a signal to channel A of the EEC where it is used to calculate the metered fuel flow. The signal is supplied to channel B internally via a CAN BUS. The fuel filter differential pressure sensor provides an output to each channel of the EEC. The fuel filter differential pressure sensor measures differential pressure across the fuel filter. A fuel filter impending bypass is detected and displayed when the differential pressure exceeds 22 psi on either engine. An L(R) ENG FUEL FILTER advisory message along with an L (R) ENG FUEL FILTER - IMPENDING BYPASS info message. A fuel filter bypass occurs at 25 psid. This is indicated by the L(R) ENG FUEL FILTER advisory message and an L (R) ENG FUEL FILTER BYPASS info message. If both fuel filters are bypassed, fuel contamination can be suspected and an L-R ENG FUEL FILTER caution message is displayed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-82
CS130_73_L3_Indicating_DetailedDescription.fm
73 - Engine Fuel and Control 73-30 Indicating
CLB
L ENG FUEL FILTER
92.2
92.2
73.0
ENG FUEL FILTER - IMPENDING BYPASS
73.3
N1
SYNOPTIC PAGE - INFO 718
722
EGT
86.5 1090 124 116
FUEL FLOW METER
86.6 1110 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
10905
TOTAL FUEL (KG) 2735
FUEL FILTER DIFFERENTIAL PRESSURE SENSOR
1400 5.7 ELEV 570
2735 0
CAB ALT
RATE
ó3
CREW OXY 1850
LDG TEMP (°C)
EEC
5435
24
TRIM NU STAB
HI
23
23
ND
4.2 NL RUDDER NR
INFO
Channel A
Channel A EICAS
Channel B
Channel B
STATUS
AIR
DOOR
ELEC
FLT CTRL
FUEL
HYD
AVIONIC
INFO
CB
TOTAL FUEL FUEL USED
15300 KG 1000 KG
To Fuel/Oil Manifold
20 °C 2900 KG
20 °C 2900 KG
FUEL FILTER
LEGEND EEC
APU
Electronic Engine Control ARINC 429 CAN BUS SYNOPTIC PAGE – FUEL
CS1_CS3_7330_001
9500 KG
Figure 39: Fuel Control Indicating Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-83
CS130_73_L3_Indicating_DetailedDescription.fm
73 - Engine Fuel and Control 73-30 Indicating
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine fuel indication.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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73-84
CS130_73_L2_Indicating_MonitoringTests.fm
73 - Engine Fuel and Control 73-30 Indicating
CAS MESSAGES Table 7: INFO Messages
Table 5: CAUTION Messages MESSAGE
MESSAGE
LOGIC
L - R ENG FUEL FILTER
Both engine fuel filters are bypassed (clogged filters), or impending, i.e. likely fuel contamination.
Table 6: ADVISORY Messages MESSAGE
LOGIC
L ENG FUEL FILTER
Left engine fuel filter is impending bypass or is bypassed.
R ENG FUEL FILTER
Right engine fuel filter is impending bypass or is bypassed.
L ENGINE FAULT
Loss of redundant or non-critical function for the left engine.
R ENGINE FAULT
Loss of redundant or non-critical function for the right engine.
L FUEL FLOW DEGRADED
FADEC provides synthesized fuel flow instead of measured fuel flow. Displayed L fuel flow accuracy is degraded.
R FUEL FLOW DEGRADED
FADEC provides synthesized fuel flow instead of measured fuel flow. Displayed R fuel flow accuracy is degraded.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
LOGIC
73 L (R) ENGINE FAULT - FUEL FILTER PRESS SNSR INOP
FADEC logic has determined that the fuel filter delta P sensor signal from channel A and/or B has failed.
73 L (R) ENG FUEL FILTER IMPENDING BYPASS
INFO message will be set by the left (right) engine when the fuel filter impending bypass is detected but there is no actual bypass.
73 L (R) ENG FUEL FILTER BYPASS
INFO message will be set by the left (right) engine when the fuel filter actual bypass is detected.
TECHNICAL TRAINING MANUAL
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73-85
CS130_73_L2_Indicating_MonitoringTests.fm
73 - Engine Fuel and Control 73-30 Indicating
Page Intentionally Left Blank
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73-86
CS130_73_ILB.fm
ATA 74-80 - Ignition and Starting
BD-500-1A10 BD-500-1A11
74 - Ignition and Starting
Table of Contents 74-11 Ignition ....................................................................74-80-2 General Description .....................................................74-80-2 Component Location ...................................................74-80-4 Ignition Exciter .......................................................74-80-4 Igniter Plugs ............................................................74-80-4 Ignition Cables ........................................................74-80-4 Component Information ...............................................74-80-6 Ignition Exciter .......................................................74-80-6 Ignition Cables .......................................................74-80-8 Igniter Plugs ............................................................74-80-8 Controls and Indications ............................................74-80-10 Engine Panel.........................................................74-80-10 Throttle Quadrant Assembly .................................74-80-10 Indications.............................................................74-80-12 Detailed Description ..................................................74-80-14 Normal Ignition for Automatic Start .......................74-80-14 Continuous Ignition ...............................................74-80-14 Monitoring and Tests .................................................74-80-18 CAS Messages .....................................................74-80-19 Practical Aspects........................................................74-80-20 Ignition System Test ............................................74-80-20 80-11 Starting .................................................................74-80-22
Component Information ............................................ 74-80-26 Air Turbine Starter................................................ 74-80-26 Starter Air Valve ................................................... 74-80-28 Controls and Indications ........................................... 74-80-30 Throttle Quadrant Assembly ................................ 74-80-30 Engine Panel........................................................ 74-80-30 Indications ............................................................ 74-80-32 Operation .................................................................. 74-80-34 Rotor Bow ............................................................ 74-80-34 Automatic Start..................................................... 74-80-36 Manual Start......................................................... 74-80-38 Engine Motoring ................................................... 74-80-38 Detailed Description ................................................. 74-80-40 Automatic Starting................................................ 74-80-40 Starter Duty Time ................................................. 74-80-40 Starter Re-Engagement ....................................... 74-80-40 Abnormal Starts ................................................... 74-80-42 Monitoring and Tests ............................................... 74-80-46 CAS Messages .................................................... 74-80-47 Practical Aspects ....................................................... 74-80-48 Starter Oil Servicing ............................................ 74-80-48 Starter Chip Collector........................................... 74-80-48
General Description ...................................................74-80-22 Automatic Start .....................................................74-80-22 Manual Start..........................................................74-80-22 Rotor Bow .............................................................74-80-22 Component Location .................................................74-80-24 Air Turbine Starter.................................................74-80-24 Starter Air Valve....................................................74-80-24 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-i
CS130_M5_74-80_IgnitionStartingTOC.fm
74 - Ignition and Starting
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7: Figure 8: Figure 9: Figure 10: Figure 11:
Ignition System.................................................74-80-3 Ignition System Component Location...............74-80-5 Ignition Exciter..................................................74-80-7 Ignition Cables and Igniter Plugs......................74-80-9 Ignition System Controls ................................74-80-11 Ignition System Indications.............................74-80-13 Ignition System Schematic .............................74-80-15 Ignition and Starting Modes............................74-80-17 Ignition System Test.......................................74-80-21 Engine Starting System..................................74-80-23 Engine Starting System Component Location ......................................74-80-25 Figure 12: Air Turbine Starter ..........................................74-80-27 Figure 13: Starter Air Valve .............................................74-80-29 Figure 14: Engine Start Controls .....................................74-80-31 Figure 15: Engine Starting Indications.............................74-80-33 Figure 16: Rotor Bow Altitude/Outside Air Temperature Graph ..................................74-80-35 Figure 17: Automatic Start...............................................74-80-37 Figure 18: Manual Start ...................................................74-80-39 Figure 19: Automatic and Manual Starting Schematic ....74-80-41 Figure 20: Abnormal Starts..............................................74-80-43 Figure 21: Fuel Depulsing ...............................................74-80-45 Figure 22: Starter Oil Servicing and Starter Chip Collector .....................................74-80-49
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74-ii
CS130_M5_74-80_IgnitionStartingLOF.fm
74 - Ignition and Starting
Page Intentionally Left Blank
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74-iii
CS130_M5_74-80_IgnitionStartingLOF.fm
IGNITION and STARTING - CHAPTER BREAKDOWN IGNITION AND STARTING CHAPTER BREAKDOWN
Ignition System
1
Starting System
2
74-80 - Ignition and Starting 74-11 Ignition
74-11 IGNITION GENERAL DESCRIPTION The engine ignition system consists of two electrically and physically independent circuits located in a single ignition exciter. Each exciter circuit is an AC-powered capacitor discharge type that supplies high voltage to an igniter plug through a shielded, high tension cable.
Automatic selection of continuous dual-ignition is provided by the EEC during critical or adverse conditions. The EEC also includes an automatic relight system, which energizes both igniters within 2 seconds when an engine flameout is detected.
Each exciter circuit is powered from the aircraft. The left engine ignition power is supplied by DC ESS BUS 1 and DC ESS BUS 3. Right engine ignition power is supplied by DC ESS BUS 2 and 3. The electronic engine control (EEC) provides the control of the ignition.
The EEC automatically selects a dual-igniter continuous ignition if any of the following conditions are true:
The ignition system is used during engine start or engine relight. The ignition system is also used any time ambient conditions require a continuous ignition to prevent the risk of flameout. During a normal automatic engine start, the EEC uses a single igniter plug to start the engine. The igniter plug selected alternates on each normal automatic engine start. The EEC commands the igniter plugs on or turns the igniter plugs off based on N2 speed.
•
Engine flameout is detected in-flight or during takeoff
•
Surge is detected in-flight or during takeoff
•
In-flight start
The automatic selection of continuous ignition occurs even if manual continuous ignition is selected.
The ignition system can also be operated in a dual-igniter low power continuous ignition mode. The continuous ignition mode operates automatically under the control of the EEC, or manually when selected from the CONT IGNITION PBA on the ENGINE panel. When the CONT IGNITION PBA is selected before engine start, the EEC enters the manual start mode and both igniter plugs are selected for engine start. In the manual start mode, the ignition is commanded on when the THROTTLE QUADRANT ASSEMBLY L(R) ENG switch is selected to the ON position, and turned off when the engine self-sustaining speed is reached. When continuous ignition is selected and the engine is already running, the igniter plugs only fire when the engine is operating at low power settings. This increases the life of the igniter plugs. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-2
CS130_74-80_L2_Ignition_General.fm
74-80 - Ignition and Starting 74-11 Ignition
MAX
MAX
IDLE
ID IDLE
L ENG ON
OFF
THROTTLE QUADRANT ASSEMBLY ENGINE CONT IGNITION
IGNITER PLUG A
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
DMC 1 and DMC 2
IGNITION EXCITER
EEC
IGNITER PLUG B ENGINE PANEL
DC ESS BUS 3
NOTE
DMC EEC
LEGEND
Left engine shown. Right engine similar.
Data Concentrator Unit Module Cabinet Electronic Engine Control
Right engine ignition exciter power DC ESS BUS 2 and DC ESS BUS 3
CS1_CS3_7400_002
DC ESS BUS 1
Figure 1: Ignition System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-3
CS130_74-80_L2_Ignition_General.fm
74-80 - Ignition and Starting 74-11 Ignition
COMPONENT LOCATION The ignition system includes the following components: •
Ignition exciter
•
Ignition cables
•
Igniter plugs
IGNITION EXCITER The ignition exciter is located on the left side of the engine fan case at the 8 o’clock position. IGNITER PLUGS The igniter plugs are installed in the diffuser case. Igniter A is located at the 4 o’clock position, and igniter B is located at the 5 o’clock position. IGNITION CABLES The ignition cables extend under the engine from the ignition exciter to the igniter plugs.
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74-80-4
CS130_74-80_L2_Ignition_ComponentLocation.fm
74-80 - Ignition and Starting 74-11 Ignition
Igniter Plugs
Ignition Cables
CS1_CS3_7400_019
Ignition Exciter
Figure 2: Ignition System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-5
CS130_74-80_L2_Ignition_ComponentLocation.fm
74-80 - Ignition and Starting 74-11 Ignition
COMPONENT INFORMATION IGNITION EXCITER The dual-channel ignition exciter supplies 5 kV to the igniter plugs. The ignition exciter has a duty cycle that varies between 1 spark per second to 3 sparks per second. The ignition exciter is cooled by the fan compartment ventilation air.
WARNING MAKE SURE THE IGNITION SYSTEM HAS NOT BEEN OPERATED FOR AT LEAST 5 MINUTES PRIOR TO REMOVING THE IGNITER PLUGS OR THEIR CABLES. IGNITION VOLTAGE CAN BE VERY HIGH AND COULD CAUSE INJURY.
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74-80-6
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
Ignition Cable B
Ignition Cable A
CS1_CS3_7400_005
Ignition Exciter
Figure 3: Ignition Exciter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-7
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
IGNITION CABLES Ignition cables are interchangeable with flexible braided steel and ceramic insulated terminals at the plug end. IGNITER PLUGS The igniter plugs extend through the diffuser case into the combustion chamber. Classified spacers are installed under a mounting boss and are used to control the immersion depth of the igniter plug. The spacer and boss are installed on the diffuser case assembly and are not removed during igniter plug replacement.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-8
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
Igniter Plug System A
A
Igniter Plug
Mounting Boss
Classified Spacer Igniter Plug System B
Ignition Cables IGNITER PLUG
CS1_CS3_7400_006
A
Figure 4: Ignition Cables and Igniter Plugs Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-9
CS130_74-80_L2_Ignition_ComponentInformation.fm
74-80 - Ignition and Starting 74-11 Ignition
CONTROLS AND INDICATIONS ENGINE PANEL The CONT IGNITION PBA is used to select manual ignition. The manual selection is indicated by the white ON legend on the pushbutton annunciator (PBA). THROTTLE QUADRANT ASSEMBLY The L(R) ENG switch signals the EEC to command the ignition ON during a manual engine start when they are selected to the ON position.
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74-80-10
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition and Starting 74-11 Ignition
ENGINE
ID D IDLE
IDLE
CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
L ENG
RE ENG NG
ENGINE PANEL
ON
OFF
OFF CS1_CS3_7400_007
THROTTLE QUADRANT ASSEMBLY
Figure 5: Ignition System Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-11
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition and Starting 74-11 Ignition
INDICATIONS An IGN icon is displayed below the exhaust gas temperature (EGT) gauges when ignition is selected. The IGN icon is green when the ignition system is commanded and the igniters are firing. The IGN icon is white when continuous ignition is selected ON, but inhibited.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-12
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition and Starting 74-11 Ignition
TO 90.8
90.8
0.0
0.0
N1
44
16
EGT
SYMBOL
COLOR
DESCRIPTION
IGN
Green
Ignition active (Firing).
IGN
White
Ignition selected but inhibited due to design constraints.
IGN N2 FF (KPH) OIL TEMP OIL PRESS FAN VIB
0.0 0 16 0 ---
CS1_CS3_7400_020
28.0 90 17 47 ---
EICAS PAGE
Figure 6: Ignition System Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-13
CS130_74-80_L2_Ignition_ControlsIndications.fm
74-80 - Ignition 74-11 Ignition
DETAILED DESCRIPTION When the engine is running, manual continuous ignition can be selected by cycling the CONT IGNITION PBA to NORMAL and back to ON.
NORMAL IGNITION FOR AUTOMATIC START The EEC channel A controls ignition system A, and EEC channel B controls ignition system B. A single igniter plug is used during an automatic start. During normal starts, the igniter plugs are commanded on at 20.4% to 23.5% N2. When the N2 reaches 49.3 to 53.3%, the EEC commands the igniter plugs off. A green IGN icon is shown on EICAS anytime the EEC activates the igniters. The icon is removed when the igniter plugs are not on. If the engine fails to start on the first attempt, the EEC dry motors the engine, and then selects both igniter plugs for a second start attempt.
When the engine is running and manual continuous ignition is selected, a white IGN icon is displayed on the engine indication and crew alerting system (EICAS). In this condition, the igniter plugs may not be firing unless the engine is at low power. The EEC monitors the burner pressure (Pb). If Pb is greater than 150 psi, the ignition is commanded off. Once the engine is running at approach idle or less, Pb is low and the ignition is commanded on. Automatic Continuous Ignition The EEC automatically selects dual-igniter plugs continuous ignition under any of the following conditions:
CONTINUOUS IGNITION
•
Engine flameout detected in-flight or during takeoff above 60 kt
After the engine has started, normal engine operation does not require the igniter plugs to sustain combustion. However, under certain conditions continuous ignition is available for use.
•
An engine surge is detected in air
•
In-flight start
Continuous ignition has two modes:
Quick-Relight Continuous Ignition
•
Manual
•
Automatic
The engine performs a quick-relight in-flight only when the L(R) ENG switch is cycled from ON to OFF, and back to the ON position.
Manual Continuous Ignition Manual continuous ignition is available when the CONT IGNITION PBA is pressed. An ON light illuminates on the PBA to indicate manual continuous ignition is available.
When the quick-relight condition occurs, the EEC commands continuous ignition until either the engine has been above idle for 30 seconds, or the L(R) ENG switch is placed in the OFF position.
During a manual engine start, continuous ignition is used and both igniter plugs fire once the L(R) ENG switches are selected to ON. The EEC commands the igniter plugs off when the N2 reaches 55%. A green IGN icon appears as long as the igniters are firing.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-14
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition 74-11 Ignition
MAX
MAX
EPC 1 DC ESS BUS 3
ID IDLE
IDLE
EEC L ENG
7.5
ENG IGN A
IGNITION EXCITER
Channel A
R ENG ENG
ON
Ignition Command
OFF
TQA
Channel B
DMC 1 and DMC 2
EICAS
ENGINE START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
System B
IGNITER PLUG B
P3
N2
ON
CDC 1-7-1 L ENG IGN B DC ESS 1 SSPC 7.5A Logic: Always On
ENGINE PANEL
LEGEND CDC DMC EEC EPC P3 TQA
NOTE Left engine shown, right engine similar.
Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Engine Control Electrical Power Center Burner Pressure Throttle Quadrant Assembly ARINC 429
CS1_CS3_7400_012
CONT IGNITION
IGNITER PLUG A
Ignition Command
IGN
Air/Gnd
System A
Figure 7: Ignition System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-15
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition 74-11 Ignition
Ignition and Starting Modes The START switch located on the ENGINE panel is in the AUTO position for normal engine starts. The L (R) ENG CRANK position is used for manual engine starts, dry motoring, and wet motoring. The CONT IGNITION PBA is used to initiate a manual engine start or select both igniters on. The L (R) ENG switches initiate an automatic start when the START switch is in AUTO. During a manual engine start, moving the L (R) ENG to ON initiates fuel and ignition. When the switch is moved to the OFF position, the engine shuts down. During an automatic engine start, only one igniter is used. The EEC alternates igniters on consecutive starts. During a manual engine start on the ground or an engine start in the air, both igniters are on. When the engine is dry or wet motored, the igniters are off. In addition to the ignition selection for engine starting, the EEC automatically selects dual-igniter ignition under any of the following conditions: •
In-flight engine start
•
Engine flameout detected in-flight or during takeoff above 60 kt
•
An engine surge is detected in air
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-16
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition 74-11 Ignition
ENGINE START
CONT IGNITION
L ENG CRANK
AUTO
R ENG CRANK
L ENG
R ENG
ON
ON
OFF FF FF
OFF FF FF
ON
CONT IGNITION NORMAL
START AUTO
L (R) ENG ON
Manual Start (Ground)
CONT IGNITION ON
START L (R) ENG CRANK
L (R) ENG ON
Auto Start (Air)
CONT IGNITION NORMAL
START AUTO
L (R) ENG ON
Dry Motoring
CONT IGNITION NORMAL
START L (R) ENG CRANK
L (R) ENG OFF
Wet Motoring
CONT IGNITION NORMAL
START L (R) ENG CRANK
L (R) ENG ON
Engine Shutdown
Alternates Igniters Every Start
Automatic Continuous Ignition • Engine Flameout on Takeoff or in Air • Engine Surge in Air
Air Turbine Starter (ATS) or Windmill Start
44
16
EGT IGN
L (R) ENG OFF
28.0
N2
0.0
EICAS PAGE WITH ENGINE IGN ICONS
CS1_CS3_7400_021
Auto Start (Ground)
Figure 8: Ignition and Starting Modes Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-17
CS130_74-80_L3_Ignition_DetailedDescription.fm
74-80 - Ignition and Starting 74-11 Ignition
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the starting and ignition system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-18
CS130_74-80_L2_Ignition_MonitoringTests.fm
74-80 - Ignition and Starting 74-11 Ignition
CAS MESSAGES Table 1: STATUS Message MESSAGE ENG CONT IGNITION ON
LOGIC At least one EEC receives a request from the flight deck switch to have continuous ignition.
Table 2: INFO Message MESSAGE 74 L(R) ENGINE FAULT - IGN REDUND LOSS
LOGIC An ignition circuit fault, or failure to light using an igniter, or a low side voltage fault has been detected by the left/right full authority digital engine control (FADEC) system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-19
CS130_74-80_L2_Ignition_MonitoringTests.fm
74-80 - Ignition and Starting 74-11 Ignition
PRACTICAL ASPECTS IGNITION SYSTEM TEST The ignition system test procedure checks for satisfactory operation of the ignition system. The following functions or units are verified with this test: •
EEC outputs to exciter
•
Exciter
•
Ignition cables (two per engine)
•
Igniter plugs (two per engine)
The engine igniters are wired to one channel each, with igniter 1 going to EEC channel A, and igniter 2 going to EEC channel B. As a consequence, the ignition system test can only verify the functionality of the igniter corresponding to the channel in control during the test. To test both igniters, the maintainer must run the test again on the alternate EEC channel. This test allows the user to audibly verify if the igniters provide a spark when commanded to do so.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-20
CS130_74-80_L2_Ignition_PracticalAspects.fm
74-80 - Ignition and Starting 74-11 Ignition
MAINT MENU EEC1-A-IGNITION TEST
1/4
*** WARNING *** IGNITERS WILL SPARK DURING THIS TEST A DRY CRANK MUST BE PERFORMED PRIOR TO RUNNING THIS TEST TO ENSURE NO ACCIDENTAL FUEL LIGHT OFF
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- PERFORM DRY CRANK AMP XX-XX-XX PRIOR TO THE TEST 3- FADEC IS POWERED
* NOTE * THERE IS A 5 MINUTE TIMER ON THE OPERATION OF THE IGNITER UPON PRESSING THE START BUTTON
PRESS TO START TEST
Write Maintenance Reports
Cancel
CS1_CS3_7321_023
PRESS THE START BUTTON BELOW TO INITIATE TEST
Figure 9: Ignition System Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-21
CS130_74-80_L2_Ignition_PracticalAspects.fm
74-80 - Ignition and Starting 80-11 Starting
80-11 STARTING GENERAL DESCRIPTION
MANUAL START
The engine starting system consists of an air turbine starter (ATS) and a starter air valve (SAV). The ATS uses bleed air to drive the high-pressure (HP) rotor to a high enough speed to ensure a satisfactory start. The bleed air is provided by the auxiliary power unit (APU), an external ground air cart, or the other operating engine. The SAV controls the flow of bleed air to the ATS. When the engine has reached a predetermined speed, the SAV closes, shutting off the bleed air to the ATS. As the engine continues to accelerate, the starter clutch automatically disengages the starter. The ATS can also be used to start the engine in flight if a windmill start can not be accomplished. Engine starting can be accomplished automatically or manually. AUTOMATIC START The automatic start mode provides the electronic engine control (EEC) with full control to sequence the SAV, ignition exciter and the fuel metering valve located in the integrated fuel pump and control (IFPC). The automatic start is initiated when the engine is not running, the START switch is in AUTO, and the L(R) ENG switch is set to ON. When the start command is received from the L(R) ENG switch, the EEC commands the SAV, ignition exciter, and the fuel-on function of the IFPC. Once the engine reaches the starter cutout speed, the EEC commands the SAV closed, and the ignition exciter off. Switching the L(R) ENG switch OFF during an automatic start shuts off the fuel, closes the SAV, and turns the ignition system off. The EEC monitors the automatic start and has the capability to abort the startup to the self-sustaining engine speed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Manual engine starting is available on the ground only. The manual start mode limits the authority of the EEC so that the starter, ignition, and fuel are controlled manually. This provides the ability to dry motor or wet motor the engine. The manual start sequence is available when the CONT IGNITION PBA is selected in the ON position. The start is initiated when the START switch is selected to the L (R) ENG CRANK position. The EEC commands the SAV open and the ATS drives the HP rotor. Once the engine is rotating, fuel and ignition exciter are commanded ON by selecting the L(R) ENG switch to ON. Once the engine reaches the starter cutout speed, the EEC commands the SAV closed, and the ignition exciter off. The manual start sequence can be terminated by selecting the START switch to the AUTO position if the L(R) ENG switch has not been selected ON, or by selecting the L(R) ENG switch to OFF. The EEC only provides fault indications for the engine indication and crew alerting system (EICAS) during manual operation. The EEC start abort function is not available. Monitoring of the start parameters is required. The start is terminated by selecting the L(R) ENG switch to OFF. ROTOR BOW To minimize the effects of a bowed rotor, a motor-to-start dry crank is incorporated into every engine start sequence. The starter air valve is opened and modulated to maintain an N2 speed between 7% and 13% for a minimum of 30 seconds before continuing with the engine start.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-22
CS130_74-80_L2_Starting_General.fm
74-80 - Ignition and Starting 80-11 Starting
DC ESS BUS 1 DC ESS BUS 3 ENGINE CONT IGNITION
L ENG
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
ON
OFF
ENGINE PANEL
THROTTLE QUADRANT ASSEMBLY
IGNITION EXCITER
DMC 1 AND DMC 2
EEC
IFPC
Air Data
DMC EEC IFPC T/M
Bleed Air Data Concentrator Unit Module Cabinet Electronic Engine Control Integrated Fuel Pump and Control ARINC 429 Torque Motor
T/M
AIR TURBINE STARTER
Bleed Air System
MAIN GEARBOX
STARTER AIR VALVE NOTE Left engine shown, right engine similar.
CS1_CS3_8000_004
LEGEND
N2 Speed
Figure 10: Engine Starting System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-23
CS130_74-80_L2_Starting_General.fm
74-80 - Ignition and Starting 80-11 Starting
COMPONENT LOCATION AIR TURBINE STARTER The air turbine starter (ATS) is mounted to the main gearbox (MGB), located on the left side of the engine below the air oil cooler. STARTER AIR VALVE The starter air valve (SAV) is mounted to the MGB, located on the left side of the engine behind the air cooler.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-24
CS130_74-80_L2_Starting_ComponentLocation.fm
74-80 - Ignition and Starting 80-11 Starting
Starter Air Valve
CS1_CS3_8000_005 00 005
Air Turbine Starter
Figure 11: Engine Starting System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-25
CS130_74-80_L2_Starting_ComponentLocation.fm
74-80 - Ignition and Starting 80-11 Starting
COMPONENT INFORMATION AIR TURBINE STARTER The starter converts power from the bleed air system to rotational energy using a turbine and a reduction gearbox. The gearbox drives the output shaft through a ratchet and pawl clutch. At low engine speeds, the pawl springs force the pawls to engage with the ratchet on the output shaft. When the engine speed exceeds that of the output hub, the pawls act as fly weights and disengages from the drive shaft. To prevent damage to the air turbine stater (ATS) or the engine during start, the output shaft has a shear section between it and the main gearbox (MGB). This is designed to fail within a specific torque range. Starter speed is sent to the EEC by a dual-channel speed sensor located on the starter housing. The speed sensor is a line replaceable unit (LRU) that provides a starter speed signal to each channel of the engine electronic control (EEC). The EEC uses the starter speed to determine start valve position, and starter engagement and disengagement. The starter has a sight glass to verify the oil level, a fill port for servicing the oil, and a starter chip collector. The starter chip collector is to prevent any debris from circulating through the ATS oil system. A locator pin on the ATS gearbox interface is used to ensure the ATS is properly clocked during installation. The ATS is cooled using fan air from the inlet of the air/oil cooler. The air flows through a cooling ring around the ATS forward flange.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-26
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
FULL
Locator Pin ADD
Output Shaft
A
SIGHT GLASS
A
Dual-Channel Speed Sensor
Starter Sight Glass Chip Collector
Fill Port
CS1_CS3_8000_006
Cooling Ring
Figure 12: Air Turbine Starter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-27
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
STARTER AIR VALVE
Manual Override Start Procedure
The starter air valve (SAV) is a pneumatically-actuated butterfly valve. A pressure regulator maintains an output pressure of 30 psi. The valve is spring-loaded to the closed position. The SAV is EEC controlled using a dual-coil torque motor. During start, the valve opening is confirmed through starter speed (Ns) and the N2.
The following procedure is used during a manual override start:
A filter port is accessible once the shroud is removed to access a line replaceable filter.
4. Ensure the SAV is in the closed position. Rotate the square drive clockwise to confirm.
A cooling shroud surrounds the SAV. A hose provides air from the air/oil cooler to the shroud to protect the valve from the heat of the engine core.
5. When directed, open the SAV by turning the manual override counterclockwise until the stop is contacted.
If the SAV fails to operate, a manual override is provided to open the valve. An SAV manual override access hole on the cowls permits a 3/8 in square drive manual override to be inserted into the SAV for manual opening.
6. Hold the valve in the open position.
1. Establish communication with the flight deck. 2. Ensure bleed air is available to the SAV. 3. Insert the 3/8 in square drive into the SAV manual access hole.
7. When directed, allow the valve to return to the closed position. The valve is spring-loaded to the closed position. Rotate the square drive clockwise to close if required. 8. Remove the manual override drive. 9. Move clear of the aircraft, avoiding the engine hazard areas. 10. Contact the flight deck to indicate clear of the aircraft.
NOTE This is not an approved procedure at this time due to the incorporation of the motor-to-start logic for rotor bow prevention.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-28
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
Cooling Shroud (4x)
Starter Air Valve Manual Override Starter Air Valve (SAV) Override Access Hole
Torque Motor
CS1 CS3 8000 008 CS1_CS3_8000_008
Filter Port
Figure 13: Starter Air Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-29
CS130_74-80_L2_Starting_ComponentInformation.fm
74-80 - Ignition and Starting 80-11 Starting
CONTROLS AND INDICATIONS THROTTLE QUADRANT ASSEMBLY The L(R) ENG switches are used to initiate an automatic start, or command the fuel and ignition during a manual start. The L(R) ENG switches are also used to shut down the engine. ENGINE PANEL The engine panel has a three-position START switch marked AUTO and L(R) ENG CRANK. The switch is spring-loaded to the AUTO position. When a start is initiated with the switch in the AUTO position, an automatic start is performed. The L(R) ENG CRANK positions are used for manual starts and wet or dry motoring. The CONT IGNITION PBA is used to select manual ignition. It also selects the manual start mode for the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-30
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
ENGINE
ID D IDLE
IDLE
CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
L ENG
ENGINE PANEL
ON
OFF
OFF CS1_CS3_8000_010
THROTTLE QUADRANT ASSEMBLY
Figure 14: Engine Start Controls Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-31
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
INDICATIONS While the EEC controls the start sequence, the engine indication and crew alerting system (EICAS) page is monitored during the start sequence. Indications that are monitored during engine start include: •
N1
•
Exhaust gas temperature (EGT)
•
Ignition
•
N2
•
Fuel flow
•
Oil pressure
The following indications are displayed above the N1 gauges, depending upon start mode: •
START - Engine is in the start sequence
•
RELIGHT - Indicates an automatic relight
In the case of an automatic restart in flight, there are the following indications depending on the type of restart: •
WINDMILL - Altitude from 30,000 ft to sea level, airspeed above 250 KIAS
•
ATS - Indicates an air turbine starter start mode. There are two different types of ATS starts -
Starter assisted APU - Altitude from 23,000 ft to sea level
-
Engine cross-bleed - Altitude from 30,000 ft to sea level, airspeed less than 250 KIAS
The EEC determines what type of in-flight restart to use depending on the altitude and airspeed of the aircraft.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-32
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
TO
START 90.8
90.8
0.0
0.0
N1 START ICONS
44
16
RELIGHT ATS
EGT IGN
28.0 90 17 47 ---
START
WINDMILL N2
FF (KPH) OIL TEMP OIL PRESS FAN VIB
0.0 0 16 0 ---
EICAS PAGE
Description Engine in start sequence. Automatic relight. ATS (Air Turbine Start). Windmill.
CS1_CS3_7400_008
Symbol
Figure 15: Engine Starting Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-33
CS130_74-80_L2_Starting_ControlsIndications.fm
74-80 - Ignition and Starting 80-11 Starting
OPERATION ROTOR BOW The motor-to-start time for rotor bow prevention is based on an altitude/outside air temperature (OAT) graph. Within the altitude/OAT envelope, the motor to start time is 30 seconds. As the distance outside the altitude/OAT envelope increases, the motor-to-start time increases. An ENG START DELAY advisory message is displayed anytime an engine start is initiated outside of the envelope. The message clears when the motor to start phase is complete.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-34
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
ALT
14000 ft
30 Second Motor-To-Start Time
5000 ft
ENG START DELAY
-2000 ft
OAT -17°C
-5°C
+43°C
CS1_CS3_7400_022
ENG START DELAY
Figure 16: Rotor Bow Altitude/Outside Air Temperature Graph Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-35
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
AUTOMATIC START When the L(R) ENG switch is moved to the ON position, the electronic engine control (EEC) begins the start sequence: •
Starter air valve (SAV) is opened
•
Oil pressure increases
•
START icon displayed
•
ENG START DELAY advisory message on if motor-to-start time > 30 seconds
•
Engine dry cranks between 7%-13% N2 for a minimum of 30 seconds
•
ENG START DELAY message off
•
At 20.4% to 23.5% N2, the EEC commands ignition using one of the igniter plugs
•
IGN icon on
•
Commands the fuel metering valve in the integrated fuel pump and control (IFPC) to supply a fuel flow of 91 kg/hr (200 lb/hr)
•
Within 20 seconds of fuel flow: -
EGT increases
-
N1 increases before 45% N2
•
At 49.3 to 53.3% N2, the EEC commands the starter air valve closed and shuts off the ignition
•
IGN icon off
•
At engine idle, 62% N2, fuel flow reaches 227 kg/hr (500 lb/hr)
•
START icon off
For automatic starts in flight, the EEC selects both igniters on.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-36
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
100
80 ENG START DELAY Advisory Message OFF
N2 (%) 40
20
SAV Modulates to Maintain 7-13% N2
0
49.3% to 53.3% N2 45% N2
• EGT rise • Fuel flow increase • N1 increase after 38% N2
30 seconds minimum
Engine IDLE FF: 227 kg/hr START Icon OFF
62% N2
20.4 to 23.5% N2 EEC commands: • Fuel ON • Ignition ON • IGN Icon ON
60
49.3% to 53.3% N2 EEC commands: • SAV closed • Ignition OFF • IGN Icon OFF
15 seconds TIME →
EEC EGT FF kg/hr SAV
Electronic Engine Control Exhaust Gas Temperature Fuel Flow Kilograms per Hour Starter Air Valve
• L(R) ENG Switch to ON • Starter Air Valve Open • START Icon On • ENG START DELAY Advisory Message on if motor-to-start time > 30 seconds • Oil Pressure Increases
L ENG ON
ENGINE CONT IGNITION
START L ENG CRANK
AUTO
R ENG CRANK
ON
OFF
THROTTLE QUADRANT ASSEMBLY
ENGINE PANEL
PILOT EVENT
NOTE Three starter crank cycles for a total of four minutes cranking time maximum.
CS1_CS3_7400_010
LEGEND
Figure 17: Automatic Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-37
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
MANUAL START In a manual start, the CONT IGNITION PBA is selected to ON. The START switch is held in the L(R) ENG CRANK position, and the following occurs; •
Starter air valve opens
•
ENG START DELAY advisory message on if motor-to-start time > 30 seconds
•
Engine dry cranks between 7%-13% N2 for a minimum of 30 seconds
•
ENGINE START DELAY message off
•
At 18.5% to engine maximum motoring speed of 24% N2, the L(R) ENG switch is moved to RUN
•
•
When the engine idle speed is reached the CONT IGNITION PBA is selected NORMAL. ENGINE MOTORING When on ground, the engine can be motored for maintenance purposes. The engine can run continually for 4 minutes without stopping. A wait time of 30 minutes is required before starting again. There are two motoring sequences, dry and wet motoring. Dry Motoring The EEC initiates a dry motoring when bleed air is available and the following controls are set: •
START switch in L(R) ENG CRANK
-
START icon on
•
CONT IGNITION PBA in normal
-
The EEC commands the ignition using both igniter plugs
•
L(R) ENG switch OFF
-
IGN icon on
-
Commands the fuel metering valve in the integrated fuel pump and control (IFPC) to open and supply a fuel flow of 91 kg/hr (200 lb/hr)
The motoring can be interrupted any time by releasing the START switch to AUTO.
The EEC initiates a wet motoring when bleed air is available and the following controls are set:
Within 20 seconds of fuel flow: -
EGT increases
-
N1 increases before 45% N2
Wet Motoring
•
Pull the applicable ignition circuit breakers or open the ignition solid-state power controllers (SSPCs)
At 49.3 to 53.3% N2, the START switch is released to the AUTO position then the EEC commands the SAV closed and shuts off the ignition
•
START switch in L(R) ENG CRANK
•
CONT IGNITION PBA in normal
-
•
L(R) ENG switch to RUN
IGN icon OFF
•
At engine idle, 62% N2, fuel flow reaches 227 kg/hr (500 lb/hr)
•
Start icon off
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Positioning the L(R) ENG switch to the OFF position shuts off fuel and initiates a dry crank. The motoring can be interrupted any time by releasing the engine START switch to AUTO.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-38
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
100
80
60
18.5 to 24% N2 • L(R) ENG Switch to RUN • START Icon On EEC commands: • Fuel ON • Ignition ON • IGN Icon ON
ENG START DELAY Advisory Message OFF
N2 (%)
Engine IDLE FF: 227 kg/hr START Icon OFF
62% N2 49.3 to 53.3% N2 45% N2
40 Oil Pressure Increasing
20
0
SAV Modulates to Maintain 7-13% N2
• EGT rise • Fuel flow increase • N1 increase after 38% N2
30 seconds minimum
49.3 to 53.3% N2 START Switch to AUTO EEC commands: • SAV closed • Ignition OFF • IGN Icon OFF
15 seconds TIME →
EEC EGT FF kg/hr SAV
L ENG ON
ENGINE CONT IGNITION
NOTE
START L ENG CRANK
AUTO
R ENG CRANK
ON
OFF F
THROTTLE QUADRANT ASSEMBLY
ENGINE PANEL
PILOT EVENT
Three starter crank cycles for a total of four minutes cranking time maximum. Four minutes maximum continuous cranking.
CS1_CS3_7400_017
• CONT IGNITION PBA ON • START Switch to L(R) ENG CRANK Electronic Engine Control • Starter Air Valve Open Exhaust Gas Temperature • ENG START DELAY Advisory Fuel Flow Message on if motor-to-start time Kilograms per Hour > 30 seconds Starter Air Valve
LEGEND
Figure 18: Manual Start Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
74-80-39
CS130_74-80_L2_Starting_Operation.fm
74-80 - Ignition and Starting 80-11 Starting
DETAILED DESCRIPTION
not reached and the START icon disappears. The message clears when the L(R) ENG switch is selected to OFF.
AUTOMATIC STARTING
The EEC start abort is inhibited above 49.3% N2 to 53.3% N2 on the ground and at all conditions in-flight. Following a ground start abort, an automatic 30 seconds engine motoring period is provided to clear fuel vapor and to cool the engine. An automatic start performs one restart attempt on the ground.
During engine start, the EEC modulates the SAV to maintain the N2 speed between 7% and 13% for a minimum of 30 seconds to prevent rotor bow. An ENG START DELAY advisory message is displayed if the motor to start time exceeds 30 seconds. During a normal start, the starter air valve (SAV) and ignition exciter are automatically turned off by the electronic engine control (EEC) between 49.3% and 53.3% N2 speed, depending on ambient conditions. A normal engine start is indicated by the START icon above the N1 indicator. Starter assist is commanded by the EEC for in-flight starts at low Mach numbers where windmilling conditions are insufficient for engine starting. The EEC selects the starter assistance mode if altitude or Mach number are not available. During an in-flight start, the EEC depressurizes the engine-driven hydraulic pump (EDP) to unload the engine and improve starting performance. A windmill start is carried out if the altitude is between 30,000 ft and sea level with airspeed above 250 KIAS. If the aircraft is not in the windmill start envelope, the starter can be used to assist the start. If the altitude is between 30,000 ft and 23,000 ft, the engine can be restarted using bleed air from the opposite engine. If the altitude and speed are below 23,000 ft and less than 250 kt, the auxiliary power unit (APU) can be used for starting. On the ground, upon detection of a no light, hot or hung start, loss of exhaust gas temperature (EGT) or a locked low-pressure rotor, the EEC automatically shuts off fuel, ignition, and starter air and provides the appropriate fault indication to the flight deck. An automatic start abort is indicated by an L(R) ENG START ABORT caution message on the engine indication and crew alerting system (EICAS). The message indicates that the engine start procedure has aborted. Idle speed was
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
STARTER DUTY TIME A STARTER TIME EXCEEDED caution message is displayed during auto or manual starts when the starter operating time reaches 4 minutes, or if the starter is operated above 44% N2 for 30 seconds. A 30 minute cooling time is required after three start attempts or 4 minutes of continuous cranking. The EEC monitors the starter operating time including during both wet and dry cranking. During an automatic start on the ground, the start attempt is discontinued and the EEC closes the starter air valve. When the message occurs, the L(R) ENG switch must be selected OFF to reset the EEC timers. For a manual start or an in-flight start, the message remains displayed until the start is terminated, or the engine reaches idle. STARTER RE-ENGAGEMENT The L(R) ENG switch must be selected OFF before a second attempt can be made following an aborted manual start attempt or an engine shutdown. If the starter engagement is interrupted during a manual start, the starter can be re-engaged below the starter cut-out speed. During manual or automatic restarts, the EEC prevents starter re-engagement until the starter rotational speed is synchronized with engine N2 speed. The starter re-engagement speed is 20% N2. The EEC compares N2 with the starter speed signal. If the starter speed is unavailable, the start valve must be closed for 1 minute, then the starter can be re-engaged at 20% N2 on the ground, or in the air.
TECHNICAL TRAINING MANUAL
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74-80-40
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
ENGINE CONT IGNITION
L ENG
START L ENG CRANK
AUTO
R ENG CRANK
PILOT EVENT
ON
ON
OFF
EDP
THROTTLE QUADRANT ASSEMBLY
ENGINE PANEL
EEC
Channel A
IGN EICAS
DMC 1 and DMC 2
IFPC
Channel B IGNITION EXCITER
ATS Air Data • Altitude • MACH
Starter Speed
ATS DMC EDP EEC IFPC
Air Turbine Starter Data Concentrator Unit Module Cabinet Engine-Driven Hydraulic Pump Electronic Engine Control Integrated Fuel Pump And Control ARINC 429
STARTER AIR VALVE
CS1_CS3_8011_001
LEGEND
Figure 19: Automatic and Manual Starting Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-41
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
ABNORMAL STARTS Normal Engine Start
Automatic Start Abort
If the oil pressure is less than 70 psi after 30 seconds at the usual idle speed, shut down the engine using the normal shutdown procedure.
The EEC aborts an automatic start if the engine encounters one or more of the following events during the start, and are not recoverable using the fuel depulse or automatic restart logic.
The engine should be operated at idle power for a minimum of 5 minutes before shutdown. If necessary, the minimum idle time can be reduced to 3 minutes. Automatic Start Failure The EEC detects an automatic start failure if there are any of the following faults detected: •
Slow N2 spool-up
•
Rapidly increasing EGT
•
N1 roll back
•
N1 locked rotor: The EEC determines no N1 engine low-pressure rotor speed after sensing N2 has increased beyond 30%
•
EEC unable to command igniters: The EEC is unable to command either igniter in either channel
•
Loss of EGT indication
•
Inability to control fuel flow
When N2 is above 50%, the automatic start abort function is not available. The L(R) ENG switch must be moved to OFF to abort the engine start.
Auto Restart On ground, if the EEC senses an automatic start failure, the EEC turns the fuel and ignition OFF, and commands a dry motor to clear fuel from the engine for 30 seconds. After motoring the engine, the EEC automatically commands an auto restart. If the auto restart attempt fails, the EEC aborts the start. If the EEC detects an overspeed or a fuel metering valve fault, the EEC does not attempt an auto restart.
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TECHNICAL TRAINING MANUAL
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74-80-42
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
Automatic Start Begins Slow N2 Spool-Up
Yes
No
Yes
Overspeed Detected
Automatic Start Abort
Fail
Auto Restart
Successful
Engine Running
No Yes
Rapid EGT Rise or Loss of Signal
Yes
No
Fuel Metering Valve Operational
No
No
Engine Dry Spool 30 seconds
Yes No
N1 Roll Back
Yes
Igniters Operational
No
Yes
No EGT Signal
Automatic Start Continues
Yes
No N1 @ 30% N2
No
CS1_CS3_7400_015
Yes
Figure 20: Abnormal Starts Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-43
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
Fuel Depulsing
Engine Start Limits
Fuel depulsing is an EEC function that uses the fuel metering valve (FMV) in the integrated fuel pump and control (IFPC) to turn fuel on and off, in an attempt to clear an impending surge or hot start.
The following table summarizes the engine starting limits.
The EEC performs fuel depulsing when a surge or an impending hot start, as indicated by a rapid rise in EGT or an exceedance of 1054°C (1929° F). If an impending hot start is detected, the EEC cycles fuel off for 2 seconds and back on for 12 seconds using the fuel metering valve. The EEC performs this cycling until EGT drops below the impending hot start detection limit or the surge has cleared. Fuel depulsing is inhibited whenever any of the following conditions are present: •
The engine N2 is above 50%
•
The aircraft is flying above an altitude of 20,000 ft
•
EGT signal is lost
Table 3: Engine Starting Limits START LIMITS Maximum EGT
1054°C (1929°F)
Minimum oil temperature for starting
-40°C (-40°F)
Minimum oil temperature for operation above idle
-6°C (21°F)
Maximum Oil Temperature Starter Duty Cycle
174°C (345°F) Three starter crank cycles for a total of 4 minutes cranking time maximum followed by 30 minutes of cooling time. 4 minutes maximum continuous cranking.
A timer in the EEC is used to discontinue fuel depulsing once the maximum allowable time of 28 seconds on the ground or 140 seconds in-flight has been reached, or if EGT increases to the starting amber value. The time limit is sufficient for two depulse cycles if the aircraft is on the ground. If the time limit has been exceeded, or the EGT has reached the starting amber value, the EEC terminates fuel depulsing and dry motors the engine for 30 seconds.
Start Abort
•
No N1 rotation by 30% N2
(EEC aborts start automatically below 54% N2)
•
N2 fails to reach stabilized idle speed within 120 seconds
•
No EGT rise after 20 seconds from positive fuel flow indication
•
Oil pressure does not increase within 30 seconds
•
EGT greater than 1054°C (Hot Start)
When the aircraft is in flight and below 20,000 ft, the timer in the EEC is changed to 140 seconds for fuel depulsing.
CS1_CS3_TB74-80_001
The EEC conducts a ground auto restart when the EGT has decreased below the impending hot start temperature, which the EEC determines depending on aircraft weight-on-wheels (WOW) status and N2 speed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-44
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
Impending Hot Start or Sub-Idle Compressor Stall N2 > 50%
Yes
Hot Start or Sub-Idle Compressor Stall Clear
Fuel Depulsing Inhibited
No
ALT > 20,000 ft
Continue Automatic Start
Yes
Dry Motor 30 seconds, then Auto Restart
Yes
Automatic start Abort and Dry Motor 30 seconds
No
Time Exceeded Automatic Start 28 seconds on Ground or 140 seconds in Flight
Fuel Depulsing Maximum 140 seconds
Yes
No
No
EGT signal is lost
Yes
Yes
Fuel Depulsing Maximum 28 seconds Yes
In Flight
No
Hot Start 1054 °C
CS1_CS3_7400_018
No
No
Figure 21: Fuel Depulsing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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74-80-45
CS130_74-80_L3_Starting_DetailedDescription.fm
74-80 - Ignition and Starting 80-11 Starting
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the starting and ignition system.
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TECHNICAL TRAINING MANUAL
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74-80-46
CS130_74-80_L2_Starting_MonitoringTests.fm
74-80 - Ignition and Starting 80-11 Starting
CAS MESSAGES
Table 6: INFO Messages Table 4: CAUTION Messages
MESSAGE
MESSAGE
80 L ENGINE INFO message will be set when one of the following is FAULT- STARTER true: SYSTEM INOP A dual channel starter speed signal fault, OR a dual channel SAV fault.
LOGIC
L ENG START ABORT
L engine starting procedure aborted.
R ENG START ABORT
R engine starting procedure aborted.
LOGIC
L ENG STARTER ATS is not disengaging or the L starter air valve (SAV) FAIL ON failed to open. R ENG STARTER ATS is not disengaging or the R starter air valve (SAV) FAIL ON failed to open.
A dual channel SAV failed closed fault has been detected by the right FADEC system. In the case of dual channel sensor fault, if the opposite channel control processor (CPU) has failed, a single channel fault would set this message.
Table 5: ADVISORY Messages MESSAGE
LOGIC
L (R) ENG FAULT Loss of redundant or non-critical function for the left or right engine. L (R) ENG STARTER OVHT
Left or right starter usage does not fulfill the duty cycle criteria: A 30 minute cool down time for 4 minutes (low N2), or 30 seconds (high N2).
ENGINE START DELAY
The engine is performing dry crank cooling during a ground start for protection against rotor bow.
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74-80-47
CS130_74-80_L2_Starting_MonitoringTests.fm
74-80 - Ignition and Starting 80-11 Starting
PRACTICAL ASPECTS STARTER OIL SERVICING The starter oil is changed at specific intervals as per the aircraft maintenance schedule. To service the oil, both the fill and overfill plugs are removed. The oil is added until it spills from the overfill plug, at which point the starter sight glass should read full. STARTER CHIP COLLECTOR The air turbine starter (ATS) chip collector can be removed by hand without draining the oil. A check valve in the fitting prevents the loss of oil when the starter chip collector is removed.
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TECHNICAL TRAINING MANUAL
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74-80-48
CS130_74-80_L2_Starting_PracticalAspects.fm
74-80 - Ignition and Starting 80-11 Starting
Fill Port
Sight Glass L FU
L
Overfill Port
Drain Port
Starter Chip Collector
CS1_CS3_7400_014
D AD
Figure 22: Starter Oil Servicing and Starter Chip Collector Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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74-80-49
CS130_74-80_L2_Starting_PracticalAspects.fm
74-80 - Ignition 80-11 Starting
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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74-80-50
CS130_74-80_ILB.fm
ATA 75 - Engine Air
BD-500-1A10 BD-500-1A11
75 - Engine Air
Table of Contents 75-24 Turbine Active Clearance Control System....................75-2 General Description .........................................................75-2 Component Location ........................................................75-4 Detailed Description .........................................................75-6 Monitoring and Tests .......................................................75-8 CAS Messages ............................................................75-9 75-25 Turbine Cooling Air ....................................................75-10 General Description .......................................................75-10 75-26 Buffer Air System........................................................75-12 General Description ........................................................75-12 Component Location ......................................................75-14 Buffer Air Heat Exchanger .........................................75-14 Buffer Air Shutoff Valve..............................................75-14 Buffer Air Shutoff Valve Solenoid...............................75-14 Buffer Air Pressure Sensor ........................................75-14 Buffer Air Check Valve...............................................75-14 Monitoring and Tests .....................................................75-16 CAS Messages ..........................................................75-17
High-Pressure Compressor Stator Vane Actuator .... 75-20 Detailed Description ...................................................... 75-22 Monitoring and Tests .................................................... 75-24 CAS Messages ......................................................... 75-25 75-32 Compressor Bleed Air System................................... 75-26 General Description ...................................................... 75-26 Low-Pressure Compressor Bleed Air System ........... 75-26 High-Pressure Compressor Bleed Air System .......... 75-26 Component Location ..................................................... 75-28 2.5 Bleed Air Valve ................................................... 75-28 2.5 Bleed Valve Actuator........................................... 75-28 High-Pressure Compressor Bleed Valve .................. 75-28 High-Pressure Compressor Bleed Valve Pressure Sensor ....................................................... 75-28 Detailed Description ..................................................... 75-30 Monitoring and Tests .................................................... 75-32 CAS Messages ......................................................... 75-33 Practical Aspects ........................................................... 75-34 Electronic Engine Control Actuator Test ................... 75-34
75-31 Variable Stator Vane System .....................................75-18 General Description .......................................................75-18 Low-Pressure Compressor ........................................75-18 High-Pressure Compressor .......................................75-18 Component Location ......................................................75-20 Low-Pressure Compressor Inlet Guide Vane Bellcrank and Synchronizing Ring .............................75-20 Low-Pressure Compressor Stator Vane Actuator......75-20 High-Pressure Stator Vane Bellcrank and Synchronizing Rings ...........................75-20 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-i
CS130_M5_75_EngineAirTOC.fm
75 - Engine Air
List of Figures Figure 1: Turbine Active Clearance Control System ............75-3 Figure 2: Turbine Active Clearance Control System Component Location .............................................75-5 Figure 3: Turbine Active Clearance Control System Schematic..............................................................75-7 Figure 4: Turbine Cooling Air..............................................75-11 Figure 5: Buffer Air System ................................................75-13 Figure 6: Buffer Air System Component Location ..............75-15 Figure 7: Variable Stator Vane System .............................75-19 Figure 8: Variable Stator Vane Control System Component Location ...........................................75-21 Figure 9: Variable Stator Vane Control Detailed Description ............................................75-23 Figure 10: Compressor Bleed Air System ............................75-27 Figure 11: Compressor Bleed Air System Component Location ...........................................75-29 Figure 12: Compressor Bleed Air System Schematic ..........75-31 Figure 13: Electronic Engine Control Actuator Test .............75-35
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75-ii
CS130_M5_75_EngineAirLOF.fm
75 - Engine Air
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TECHNICAL TRAINING MANUAL
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75-iii
CS130_M5_75_EngineAirLOF.fm
ENGINE AIR SYSTEMS - CHAPTER BREAKDOWN ENGINE AIR CHAPTER BREAKDOWN
Engine Cooling
1
Compressor Airflow
2
75 - Engine Air 75-24 Turbine Active Clearance Control System
75-24 TURBINE ACTIVE CLEARANCE CONTROL SYSTEM GENERAL DESCRIPTION The turbine active clearance control system provides fan air to the turbine cases to limit turbine case growth due to thermal expansion. This reduces high-pressure turbine (HPT) and low-pressure turbine (LPT) blade tip clearance and improves fuel efficiency. The turbine active clearance control (ACC) system consists of a fuel actuated turbine ACC air valve that regulates the fan airflow to the LPT and HPT cooling air manifolds. The cooling air manifolds spray air on the turbine case. The turbine ACC air valve is controlled by the electronic engine control (EEC). Servo fuel, supplied from the integrated fuel pump and control (IFPC) positions the valve in response to EEC commands.
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75-2
CS130_75_L2_TACCS_General.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
ELECTRONIC ENGINE CONTROL
Fan Air
TURBINE ACTIVE CLEARANCE CONTROL AIR VALVE
HIGH-PRESSURE TURBINE MANIFOLD
LEGEND
LOW-PRESSURE TURBINE MANIFOLD
INTEGRATED FUEL PUMP AND CONTROL
CS1_CS3_7524_005
Fan Air Servo Fuel Servo Fuel Return
Figure 1: Turbine Active Clearance Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-3
CS130_75_L2_TACCS_General.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
COMPONENT LOCATION The turbine active clearance control system has the following components: •
Turbine active clearance control air valve
•
High-pressure turbine (HPT) manifold
•
Low-pressure turbine (LPT) manifold
Turbine Active Clearance Control Air Valve The turbine active clearance control (ACC) air valve is located at the 1 o’clock position, near the rear of the high-pressure compressor (HPC). High-Pressure Turbine Manifold The high-pressure turbine (HPT) manifold is located on the HPT case. Low-Pressure Turbine Manifold The low-pressure turbine (LPT) manifold is located on the LPT case.
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75-4
CS130_75_L2_TACCS_ComponentLocation.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
LPT Manifold
Turbine Active Clearance Control Air Valve
CS1_CS3_7524_001
HPT Manifold
LEGEND HPT LPT
High-Pressure Turbine Low-Pressure Turbine
Figure 2: Turbine Active Clearance Control System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-5
CS130_75_L2_TACCS_ComponentLocation.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
DETAILED DESCRIPTION The turbine ACC air valve has a torque motor (T/M) that positions the valve using IFPC servo fuel pressure. The torque motor control is done by the EEC channel in control. The turbine ACC air valve single channel linear variable differential transformer (LVDT) sends position feedback to channel A of the EEC. The EEC controls the valve according to engine parameters and altitude in combination with the turbine ACC air valve position feedback. The turbine ACC air valve operation is scheduled as a function of N2 and altitude. The turbine ACC air valve is closed during takeoff and most of climb. It can open during the end of the climb phase and is always open in cruise. The failsafe mode of the turbine ACC air valve is the closed position.
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75-6
CS130_75_L3_TACCS_DetailedDescription.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
PMAG N2 SPEED
DMC 1 and DMC 2
Air Data
EEC Channel B
IFPC
Channel A
TURBINE ACC AIR VALVE T/M
LVDT
HPT MANIFOLD
Fan Bypass Air LEGEND Data Concentrator Unit Module Cabinet Electronic Engine Control High-Pressure Turbine Integrated Fuel Pump and Control Low-Pressure Turbine Linear Variable Differential Transformer Permanent Magnet Alternator Generator Torque Motor Fan Air Servo Fuel Servo Fuel Return
LPT MANIFOLD CS1_CS3_7524_004
DMC EEC HPT IFPC LPT LVDT PMAG TM
Figure 3: Turbine Active Clearance Control System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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75-7
CS130_75_L3_TACCS_DetailedDescription.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the active clearance control system.
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TECHNICAL TRAINING MANUAL
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75-8
CS130_75_L2_TACCS_MonitoringTests.fm
75 - Engine Air 75-24 Turbine Active Clearance Control System
CAS MESSAGES Table 1: ADVISORY Message MESSAGE L (R) ENG FAULT
LOGIC Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 2: INFO Messages MESSAGE
LOGIC
75 L ENGINE FAULT - ACC FAIL CLSD
INFO message is set when one of the following is true: • A dual-channel ACC torque motor fault exists • Single channel ACC feedback fault has been detected by the left FADEC system • The turbine ACC air valve stuck closed
75 R ENGINE FAULT - ACC FAIL CLSD
INFO message is set when one of the following is true: • A dual-channel ACC torque motor fault exists • Single channel ACC feedback fault has been detected by the right FADEC system • The turbine ACC air valve stuck closed
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75-9
CS130_75_L2_TACCS_MonitoringTests.fm
75 - Engine Air 75-25 Turbine Cooling Air
75-25 TURBINE COOLING AIR GENERAL DESCRIPTION The turbine cooling air system consists of a series of tubes that provide 4th or 6th stage high-pressure compressor (HPC) air to cool parts of the engine. The 6th stage HPC bleed air cools the following engine parts; •
High-pressure turbine (HPT) 2nd stage vanes
•
HPT 2nd stage blade attachment
The 4th stage HPC bleed air cools the following parts; •
Turbine intermediate case (TIC) fairing
•
Low-pressure turbine (LPT) case
•
LPT rotor and blade attachment
There are four tubes that feed cooling air from the 6th stage HPC to the HPT 2nd stage vanes and blade attachments. The air is metered by plates between the tubes and the HPT case. The cooling air then flows through hollow vanes and exits out the trailing edge. Seven TIC cooling air tubes cool the TIC fairings, inner walls and outer walls known as transition ducts. Four are supply tubes, and three are jumper tubes. The three jumper tubes feed air through the TIC fairings to the LPT rotors and blade attachment. Three additional jumper tubes feed air into the space between the LPT case and 2nd stage vane.
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75-10
CS130_75_L2_TCA_General.fm
75 - Engine Air 75-25 Turbine Cooling Air
HPT LPT
CS1_CS3_7525_001
LEGEND 6th Stage Air to HPT 4th Stage Air to LPT High-Pressure Turbine Low-Pressure Turbine
Figure 4: Turbine Cooling Air Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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75-11
CS130_75_L2_TCA_General.fm
75 - Engine Air 75-26 Buffer Air System
75-26 BUFFER AIR SYSTEM GENERAL DESCRIPTION The engine bearing compartments are cooled using a split buffer bearing cooling system. This supplies the bearing compartments with cooling and pressurizing air from the 2.5 low-pressure compressor (LPC) exit air, or from the 4th stage high-pressure compressor air (HPC), depending on engine operating conditions. The system is operated by the controlling electronic engine control (EEC) channel. The EEC controls the system through the buffer air valve solenoid (BAVS). The BAVS opens the buffer air shutoff valve (BASOV) when the BAVS is energized. The BASOV controls the flow of 4th stage air into the low-pressure side of the buffer air system to maintain the required seal pressure in each of the engine's bearing compartments during low power settings.
At low power, the 2.5 air does not have sufficient pressure and flow to supply the buffer air system. The EEC energizes the BAVS to open the BASOV. The BASOV allows 4th stage air to supply the front bearing compartments, consisting of the no. 1, no 1.5 and no. 2 bearings, no. 3, and no. 5/6 compartments, with air for cooling and pressurizing the seals. The no. 4 bearing compartment is always supplied with cooled 4th stage air regardless of the engine’s power settings. At high power, the 2.5 air is sufficient to supply the buffer air system. The high power switchover occurs at 75% N1. The EEC de-energizes the BAVS to close the BASOV. When the BASOV closes, the BACV opens, allowing 2.5 air to supply the front bearing compartment no.3 and compartments no.5/6 with air for cooling and pressurizing the seals.
A buffer air pressure sensor (BAPS) monitors the pressure downstream of the BASOV to confirm that the BASOV is in the correct position for the engine power setting. The BAPS provides a signal to each EEC channel. The buffer air check valve (BACV) allows 2.5 air to supply the buffer air system at high power settings when the BASOV is closed. The BACV closes when the BASOV is open and supplies 4th stage air to the buffer air system. When the BACV is closed it prevents backflow from the 4th stage into the engine. The buffer air heat exchanger (BAHX) cools stage 4 air used for the no. 4 bearing compartment cooling and pressurization. The BAHX uses 2.5 air to cool the 4th stage air.
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75-12
CS130_75_L2_BAS_General.fm
75 - Engine Air 75-26 Buffer Air System
N1 SPEED SENSOR
BAHX
EEC BAVS
BASOV
SOL
BACV
No. 3 Bearing
BAPS
No. 5 Bearing
No. 6 Bearing
No. 5 and No. 6 Bearing Compartment
Forward Bearing Compartment
No. 3 Bearing Compartment
No. 4 Bearing Compartment
NOTE No. 1 Bearing
No. 1.5 Bearing
No. 2 Bearing
Engine shown at high power.
BACV BAHX BAPS BASOV BAVS EEC
2.5 Bleed Air 4th Stage Air Buffer Air Check Valve Buffer Air Heat Exchanger Buffer Air Pressure Sensor Buffer Air Shutoff Valve Buffer Air Valve Solenoid Electronic Engine Control
CS1_CS3_7526_001
LEGEND
Figure 5: Buffer Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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75-13
CS130_75_L2_BAS_General.fm
75 - Engine Air 75-26 Buffer Air System
COMPONENT LOCATION The buffer air cooling system consists of the following components: •
Buffer air heat exchanger (BAHX)
•
Buffer air shutoff valve (BASOV)
•
Buffer air shutoff valve solenoid (BAVS)
•
Buffer air pressure sensor (BAPS)
•
Buffer air check valve (BACV)
BUFFER AIR HEAT EXCHANGER The buffer air heat exchanger (BAHX) is located on the fan intermediate case in the 2.5 bleed cavity at the 12:00 o’clock position. BUFFER AIR SHUTOFF VALVE The buffer air shutoff valve (BASOV) is located on the rear of the high-pressure compressor (HPC) at the 11 o’clock position. BUFFER AIR SHUTOFF VALVE SOLENOID The buffer air shutoff valve solenoid (BAVS) is located on the right side of the HPC case at the 1 o’clock position. BUFFER AIR PRESSURE SENSOR The buffer air pressure sensor (BAPS) is located below the BAHX at the 2 o’clock position. BUFFER AIR CHECK VALVE The buffer air check valve (BACV) is located at the 12 o’clock position on the fan intermediate case.
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TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-14
CS130_75_L2_BAS_ComponentLocation.fm
75 - Engine Air 75-26 Buffer Air System
Buffer Air Heat Exchanger Buffer Air Check Valve
Gasket
A B C
A
C BUFFER AIR SHUTOFF VALVE SOLENOID
B
BUFFER AIR PRESSURE SENSOR
BUFFER AIR D SHUTOFF VALVE
CS1_CS3_7526_002
D
Figure 6: Buffer Air System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-15
CS130_75_L2_BAS_ComponentLocation.fm
75 - Engine Air 75-26 Buffer Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the buffer air system operations.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-16
CS130_75_L2_BAS_MonitoringTests.fm
75 - Engine Air 75-26 Buffer Air System
CAS MESSAGES Table 3: ADVISORY Message MESSAGE
LOGIC
L (R) ENG FAULT Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 4: INFO Messages MESSAGE
LOGIC
75 L (R) ENGINE FAULT - BAV INOP
INFO message is set by the left (right) engine when the buffer air shutoff valve is failed closed.
75 L (R) ENGINE FAULT - BACV FAIL CLSD
INFO message is set by the left (right) engine when the buffer air check valve is failed closed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-17
CS130_75_L2_BAS_MonitoringTests.fm
75 - Engine Air 75-31 Variable Stator Vane System
75-31 VARIABLE STATOR VANE SYSTEM GENERAL DESCRIPTION The variable stator vane system adjusts the position of the low-pressure compressor (LPC) inlet guide vanes and the high-pressure compressor (HPC) variable stator vanes during engine operation. This optimizes engine performance. LOW-PRESSURE COMPRESSOR
The HPC stator vane actuator moves the HPC inlet guide vane and the stator vanes of the first three stages mechanically through a bellcrank and synchronizing ring. The HPC stator vanes move between the open and closed position based on N2 speed. The HPC variable stator vanes are closed during engine startup and engine idle, and are fully open at takeoff power. The actuators require no rigging during installation. They are designed to fail in the open position to allow maximum airflow through the compressors.
The electronic engine control (EEC) controls the LPC stator vane actuator position through a torque motor. This controls a servovalve mounted on the actuator. A linear variable differential transformer (LVDT) provides feedback to the EEC. The LPC stator vane actuator is fuel actuated using servo fuel supplied by the integrated fuel pump and control (IFPC). The LPC stator vane actuator moves the LPC inlet guide vane mechanically through a bellcrank and synchronizing ring. The LPC inlet guide vane moves between the open and closed position, based on N1 speed. The LPC inlet guide vanes move toward the closed position during engine start up and idle, modulates during transient operations, and are open above idle power. HIGH-PRESSURE COMPRESSOR The EEC controls the HPC stator vane actuator position through a torque motor. This controls a servovalve mounted on the actuator. A LVDT provides feedback to the EEC. The HPC stator vane actuator is fuel actuated using servo fuel, supplied by the IFPC.
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75-18
CS130_75_L2_VSVS_General.fm
75 - Engine Air 75-31 Variable Stator Vane System
N1 SPEED SENSOR
LPC SVA
EEC
IFPC
PMAG N2 SPEED
HPC SVA
BELLCRANK and SYNCHRONIZING RING
LPC INLET GUIDE VANES
BELLCRANK and SYNCHRONIZING RINGS
HPC INLET GUIDE VANE and STATOR VANES
LEGEND
CS1_CS3_7531_006
EEC HPC IFPC LPC PMAG SVA
Servo Fuel Servo Fuel Return Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Control Low-Pressure Compressor Permanent Magnet Alternator Generator Stator Vane Actuator Mechanical Link
Figure 7: Variable Stator Vane System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-19
CS130_75_L2_VSVS_General.fm
75 - Engine Air 75-31 Variable Stator Vane System
COMPONENT LOCATION The variable stator vane control system consists of the following components: •
LPC inlet guide vane bellcrank and synchronizing ring
•
LPC stator vane actuator
•
HPC stator vane bellcrank and synchronizing rings
•
HPC stator vane actuator
LOW-PRESSURE COMPRESSOR INLET GUIDE VANE BELLCRANK AND SYNCHRONIZING RING The LPC inlet guide vane (IGV) bellcrank and synchronizing ring is mounted around the engine at the inlet of the LPC. LOW-PRESSURE COMPRESSOR STATOR VANE ACTUATOR The LPC SVA is mounted at the 5 o’clock position on the compressor intermediate case (CIC) firewall. HIGH-PRESSURE STATOR VANE BELLCRANK AND SYNCHRONIZING RINGS The HPC stator vane bellcrank and synchronizing rings are mounted on the HPC. HIGH-PRESSURE COMPRESSOR STATOR VANE ACTUATOR The HPC stator vane actuator is mounted on the HPC case on the right side of the engine at the 4 o’clock position.
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TECHNICAL TRAINING MANUAL
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75-20
CS130_75_L2_VSVS_ComponentLocation.fm
75 - Engine Air 75-31 Variable Stator Vane System
Low Pressure Compressor Synchronizing Ring
High Pressure Compressor Bellcrank High Pressure Compressor Starter Vane Actuator Low Pressure Compressor Starter Vane Actuator
Low Pressure Compressor Bellcrank
CS1_CS3_7531_001
High Pressure Compressor Synchronizing Rings
Figure 8: Variable Stator Vane Control System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-21
CS130_75_L2_VSVS_ComponentLocation.fm
75 - Engine Air 75-31 Variable Stator Vane System
DETAILED DESCRIPTION Either channel of the EEC controls the LPC SVA through a dual-coil torque motor which controls an internal servovalve. The position of the LPC SVA is sensed by a dual-coil linear variable differential transformer (LVDT). The LVDT is used to transmit the piston position to each electronic engine control (EEC) channel. If there is a power failure to the LPC SVA, the actuator positions the inlet guide vanes in the open position for maximum airflow through the LPC. The LPC SVA is positioned based on the N1 speed sensor signal. Either channel of the EEC control the HPC SVA through a dual-coil torque motor that controls an internal servovalve. The position of the HPC SVA is sensed by a dual-coil LVDT. The LVDT is used to transmit the piston position to each EEC channel. If there is a power failure to the LPC SVA, the actuator positions the inlet guide vanes and stator vanes in the open position for maximum airflow through the HPC. The HPC SVA is positioned based on the N2 speed sensing information received from the permanent magnet alternator generator (PMAG).
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75-22
CS130_75_L3_VSVS_DetailedDescription.fm
75 - Engine Air 75-31 Variable Stator Vane System
N1 SPEED SENSOR
LPC STATOR VANE ACTUATOR
LPC
TORQUE MOTOR
BELLCRANK and SYNCHRONIZING RING
Channel A Channel B
Inlet Guide Vanes
LVDT LEGEND
Channel A EEC HPC IFPC LPC LVDT PMAG
Channel B EEC Channel A
IFPC
Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Control Low-Pressure Compressor Linear Variable Differential Transformer Permanent Magnet Alternator Generator Mechanical Link Servo Fuel Servo Fuel Return
Channel B HPC HPC STATOR VANE ACTUATOR
Inlet Guide Vanes
TORQUE MOTOR Channel A
LVDT
BELLCRANK and SYNCHRONIZING RINGS
Channel A PMAG N2 SPEED
1st Stage Stator Vanes
2nd Stage Stator Vanes
Channel B 3rd Stage Stator Vanes
CS1_CS3_7531_005
Channel B
Figure 9: Variable Stator Vane Control Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-23
CS130_75_L3_VSVS_DetailedDescription.fm
75 - Engine Air 75-31 Variable Stator Vane System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for air system operations.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
75-24
CS130_75_L2_VSVS_MonitoringTests.fm
75 - Engine Air 75-31 Variable Stator Vane System
CAS MESSAGES Table 5: CAUTION Messages MESSAGE
LOGIC
L ENG OPER DEGRADED
Uncertainty on T2/TAT (degraded or default value), or HPC stator vane actuator not tracking as it should.
R ENG OPER DEGRADED
Uncertainty on T2/TAT (degraded or default value), or HPC stator vane actuator not tracking as it should.
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TECHNICAL TRAINING MANUAL
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75-25
CS130_75_L2_VSVS_MonitoringTests.fm
75 - Engine Air 75-32 Compressor Bleed Air System
75-32 COMPRESSOR BLEED AIR SYSTEM GENERAL DESCRIPTION The engine has a compressor bleed air system (BAS) to provide greater compressor stability during starting and engine transient operation. Air is bled from the rear of the low-pressure compressor (LPC) at station 2.5, and from one 6th stage bleed port in the high-pressure compressor (HPC).
In the event of an engine surge, the EEC:
LOW-PRESSURE COMPRESSOR BLEED AIR SYSTEM
When the surge detection signal clears, bleed valve and the variable stator vanes are returned to their normal positions, and continuous ignition is turned off after a timer expires.
The low-pressure compressor BAS uses a bleed valve. It is located at the exit of the LPC at station 2.5, and provides an improved surge margin during starting, low power, and transient operation. The bleed valve forms a ring around the LPC case. The bleed valve extracts dirt, rain, and ice during ground operation. Air from the bleed valve is vented to the fan air stream. The 2.5 bleed actuator has an integral torque motor assembly that controls servo fuel pressure from the integrated fuel pump and control (IFPC) to position the actuator. The 2.5 bleed valve is attached to the LPC just forward of the compressor intermediate case (CIC). The actuator acts on the bellcrank to open and close the valve. The valve is fully modulated through the electronic engine control (EEC) commands. The 2.5 bleed valve actuator is scheduled based on N1 speed and air data information. The 2.5 bleed valve actuator moves the 2.5 bleed valve ring to the open position during startup and idle, modulates during transient operations, and closes at takeoff. The failsafe position of this bleed is in the closed position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
•
Opens the bleed valve
•
Activates continuous ignition on both igniters
•
Resets the variable stator vanes to recover from the surge condition.
HIGH-PRESSURE COMPRESSOR BLEED AIR SYSTEM The high-pressure compressor (HPC) BAS discharges the 6th stage compressor air into the engine core area. The HPC bleed valve is augmented by the cowl anti-ice valves (CAIVs), which open during engine start and discharges air through the cowl anti-ice (CAI) ducting. For more information on the cowl anti-ice system (CAIS), refer to ATA 30 Ice and Rain Protection. The HPC bleed valve is a two-position poppet-type valve. The valve is spring-loaded to the open position during engine starting. The valve is closed using station 2.9 HPC air. The valve closes when HPC pressure is high enough to overcome the spring-force, and closes fully at idle power. The HPC bleed valve sensor provides the EEC with feedback on the position of the HPC bleed valve.
TECHNICAL TRAINING MANUAL
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75-26
CS130_75_L2_CBAS_General.fm
75 - Engine Air 75-32 Compressor Bleed Air System
HPC BLEED VALVE
HPC BLEED VALVE PRESSURE SENSOR
CAI VALVES
IFPC
2.5 BLEED VALVE
2.5 BLEED ACTUATOR
EEC
PMAG N2 SPEED
Air Data
CAI EEC HPC IFPC PMAG
Bleed Air Servo Fuel Servo Fuel Return Cowl Anti-Ice Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Control Permanent Magnet Alternator Generator Mechanical Link
NOTE 2.5 bleed valve position is based on N1 speed.
CS1_CS3_7532_006
LEGEND
Figure 10: Compressor Bleed Air System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-27
CS130_75_L2_CBAS_General.fm
75 - Engine Air 75-32 Compressor Bleed Air System
COMPONENT LOCATION The compressor bleed air system consists of the following components: •
2.5 bleed air valve
•
2.5 bleed valve actuator
•
High-pressure compressor (HPC) bleed valve
•
High-pressure compressor (HPC) bleed valve pressure sensor
2.5 BLEED AIR VALVE The 2.5 bleed air valve is located between the low-pressure compressor (LPC) and the high-pressure compressor (HPC). 2.5 BLEED VALVE ACTUATOR The 2.5 bleed valve actuator is located on the rear of the compressor intermediate case fireseal at the 3 o’clock position. HIGH-PRESSURE COMPRESSOR BLEED VALVE The high-pressure compressor (HPC) bleed valve is located on the HPC case at the 1 o’clock position. HIGH-PRESSURE COMPRESSOR BLEED VALVE PRESSURE SENSOR The high-pressure compressor (HPC) bleed valve pressure sensor is located at the 2 o’clock position on the compressor intermediate case fireseal.
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75-28
CS130_75_L2_CBAS_ComponentLocation.fm
75 - Engine Air 75-32 Compressor Bleed Air System
2.5 Bleed Valve Actuator
2.5 Bleed Valve
CS1_CS3_7532_001
High Pressure Compressor Bleed Valve
High Pressure Compressor Bleed Valve Pressure Sensor
Figure 11: Compressor Bleed Air System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-29
CS130_75_L2_CBAS_ComponentLocation.fm
75 - Engine Air 75-32 Compressor Bleed Air System
DETAILED DESCRIPTION A dual-channel LVDT, mounted in the valve actuator, provides feedback to the EEC proportional to the position of the piston. The EEC directs the fuel pressure supplied through a dual-coil torque servomotor system. This drives the valve to the required position. Scheduling of the 2.5 bleed valve actuator is based on N1 speed. For steady-state operation, the bleed valve is modulated between the fully opened and fully closed positions. The bleed valve position is scheduled as a function of N1 biased with altitude and Mach number. At takeoff power conditions, the bleed valve is fully closed. For starting and accelerations, the bleed valve follows the steady-state schedule. For decelerations, the bleed valve opens to protect LPC stability. In the event of an engine surge, the bleed valve opens to enhance recovery.
The HPC bleed valve pressure sensor sends a signal to the EEC to determine if the HPC bleed valve has failed in the open position, exposing the engine core to hot compressor air after the engine start. The sensor is also used by the EEC to determine if the HPC bleed valve has failed to close during engine start. If the valve is stuck closed during the engine start, the HPC could be overloaded, which could lead to an engine stall.
Upon detection of a surge during takeoff, the EEC assumes foreign object debris (FOD) and the bleed valve is partially opened to allow a 15% bleed flow. Assuming there is no other damage, this results in normal engine parameters except for a higher exhaust gas temperature (EGT) in the range of 45°C. The bleed valve remains in this position until a power setting below approach idle is selected, or the aircraft is not in the takeoff configuration. When the surge clears, the 2.5 bleed is positioned based on the normal control scheduling, and EGT is restored to normal levels. During engine start, the HPC bleed valve is open to unload the HPC. As the engine speed increases, the HPC moves to the closed position. In addition to the HPC bleed valve, the EEC also commands the cowl anti-ice valves (CAIV) open during start.
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TECHNICAL TRAINING MANUAL
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75-30
CS130_75_L3_CBAS_DetailedDescription.fm
75 - Engine Air 75-32 Compressor Bleed Air System
IFPC CAI VALVES
Torque Motor N1 SPEED SENSOR
Channel A
EEC
2.5 BLEED VALVE
Channel B Channel A 2.5 BLEED ACTUATOR
Air Data • Altitude • Mach Number
BELLCRANK
Channel B LVDT Channel A
Channel B LEGEND HPC PRESSURE SENSOR
Channel A HPC BLEED VALVE Channel B
NOTE Left engine shown. Right engine similar.
CS1_CS3_7532_005
CAI EEC HPC IFPC LVDT
HPC Bleed Air Servo Fuel Servo Fuel Return Cowl Anti-Ice Electronic Engine Control High-Pressure Compressor Integrated Fuel Pump and Controller Linear Variable Differential Transformer Mechanical Link
Figure 12: Compressor Bleed Air System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-31
CS130_75_L3_CBAS_DetailedDescription.fm
75 - Engine Air 75-32 Compressor Bleed Air System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the air system operations.
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TECHNICAL TRAINING MANUAL
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75-32
CS130_75_L2_CBAS_MonitoringTests.fm
75 - Engine Air 75-32 Compressor Bleed Air System
CAS MESSAGES Table 6: CAUTION Messages MESSAGE
LOGIC
L ENG NACELLE OVHT
Left burst duct or L HPC valve failed open, and automation failed to close it.
R ENG NACELLE OVHT
Right burst duct or R HPC valve failed open, and automation failed to close it.
Table 7: ADVISORY Message MESSAGE L (R) ENG FAULT
LOGIC Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 8: INFO Message MESSAGE 75 L (R) ENGINE FAULT - HPC BLEED VLV INP
LOGIC INFO message is set by the left (right) engine when the EEC active channel has detected that the HPC bleed valve has failed to close.
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75-33
CS130_75_L2_CBAS_MonitoringTests.fm
75 - Engine Air 75-32 Compressor Bleed Air System
PRACTICAL ASPECTS ELECTRONIC ENGINE CONTROL ACTUATOR TEST The actuator test detects faults on any of the fuel controlled actuators. The test requires the engine to be dry motored. The electronic engine control (EEC) performs interactive tests on the following components: •
The fuel metering valve in the IFPC
•
The low-pressure compressor (LPC) variable stator vane actuator
•
The high-pressure compressor (HPC) variable stator vane actuator
•
The low-pressure compressor (LPC) 2.5 bleed valve actuator
•
The turbine active clearance control (ACC) actuator
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75-34
CS130_75_L2_CBAS_PracticalAspects.fm
75 - Engine Air 75-32 Compressor Bleed Air System
MAINT MENU EEC1-A-ACTUATOR TEST
1/5
*** WARNING *** ENGINE WILL DRY CRANK / DRY MOTOR AND COMPONENTS WILL MOVE DURING TEST EDP WILL BE PRESSURIZED AND PMAG WILL BE ENERGIZED
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- FADEC POWER IS APPLIED 3- FUEL IS PRESENT IN ENGINE 4- AIR PRESSURE IS AVAILABLE AT STARTER INLET
* NOTE *
PRESS "CONTINUE" TO GO TO START PAGE
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7321_019
TEST DURATION: APPROXIMATELY 150 SECONDS
Figure 13: Electronic Engine Control Actuator Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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75-35
CS130_75_L2_CBAS_PracticalAspects.fm
75 - Engine Air 75-32 Compressor Bleed Air System
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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75-36
CS130_75_ILB.fm
ATA 76 - Engine Controls
BD-500-1A10 BD-500-1A11
76 - Engine Controls
Table of Contents 76-10 Engine Control..............................................................76-2 General Description ..........................................................76-2 Component Location ........................................................76-4 Component Information ....................................................76-6 Throttle Quadrant Assembly ........................................76-6 Detailed Description .........................................................76-8 Engine Control .............................................................76-8 Thrust Reverser Control...............................................76-8 Autothrottle Control ......................................................76-8 Thrust Lever Warning ..................................................76-8 Monitoring and Tests ......................................................76-10 CAS Messages ..........................................................76-11
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76-i
CS130_M5_76_EngineCtlsTOC.fm
76 - Engine Controls
List of Figures Figure 1: Engine Control System..........................................76-3 Figure 2: Throttle Quadrant Assembly Component Location .............................................76-5 Figure 3: Throttle Quadrant Assembly..................................76-7 Figure 4: Engine Control System Schematic........................76-9
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76-ii
CS130_M5_76_EngineCtlsLOF.fm
76 - Engine Controls
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76-iii
CS130_M5_76_EngineCtlsLOF.fm
ENGINE CONTROLS - CHAPTER BREAKDOWN ENGINE CONTROLS CHAPTER BREAKDOWN
Throttle Quadrant Assembly
1
76 - Engine Controls 76-10 Engine Control
76-10 ENGINE CONTROL GENERAL DESCRIPTION Engine control is carried out using thrust levers that are located on the throttle quadrant assembly (TQA). Forward and reverse thrust control for each engine is provided by a single lever. Engine RUN switches, also located on the TQA, are used for commanding engine starts. The left and right lever systems of the TQA are completely independent of each other. Each thrust lever supplies thrust command signals to the electronic engine control (EEC). Thrust lever position information is sent to the autothrottle system in the data concentrator unit module cabinets (DMCs), through the brake data concentrator unit (BDCU). Thrust command signals are used by the EEC to calculate the fuel scheduling required to achieve the desired thrust. The EEC then outputs a command to the integrated fuel pump and control (IFPC) to schedule the fuel flow. When starting the engine, the engine RUN switches on the TQA are moved to the RUN position. A signal is sent to the EEC to command the starter air valve (SAV), IFPC, and ignition system in the start sequence. Engine shutdown is initiated by moving the same engine RUN switch to OFF. The thrust levers are either controlled automatically by the autothrottle system, or manually. Thrust reversers (T/Rs) are deployed by moving the thrust levers into the reverse thrust range. Thrust reverser stow/deploy switches in the TQA detect that thrust reverse has been selected and supply signals to unlock the track lock units.
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76-2
CS130_76_L2_EngineControl_General.fm
76 - Engine Controls 76-10 Engine Control
THROTTLE QUADRANT ASSEMBLY IGNITION EXCITER BOX DMC MAX
MAX
IDLE
IDLE
Autothrottle
ELECTRONIC ENGINE CONTROL
INTEGRATED FUEL PUMP AND CONTROL
BRAKE DATA CONCENTRATOR UNIT L ENG
R ENG
OFF
OFF
ON
ON
STARTER AIR VALVE
THRUST REVERSER TRACK LOCK UNIT
DMC
CS1_CS3_7610_001
LEGEND Data Concentrator Unit Module Cabinet ARINC 429
Figure 1: Engine Control System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-3
CS130_76_L2_EngineControl_General.fm
76 - Engine Controls 76-10 Engine Control
COMPONENT LOCATION The engine throttle quadrant assembly (TQA) is located on the fight deck center pedestal.
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76-4
CS130_76_L2_EngineControl_ComponentLocation.fm
76 - Engine Controls
MAX
MAX
IDLE
IDLE
REV
REV
MAX REV
MAX REV
L ENG
R ENG
ON
ON
OFF
OFF
CS1_CS3_7600_015
76-10 Engine Control
Figure 2: Throttle Quadrant Assembly Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-5
CS130_76_L2_EngineControl_ComponentLocation.fm
76 - Engine Controls 76-10 Engine Control
COMPONENT INFORMATION
Rotary Variable Differential Transformers
THROTTLE QUADRANT ASSEMBLY
There are two types of rotary variable differential transformers (RVDTs) in the TQA, dual and single channel. Each thrust lever is connected to one of each through mechanical gears inside the TQA.
The throttle quadrant assembly (TQA) is a line replaceable unit (LRU) that houses both engine thrust levers and the L(R) ENG switches. The components that are part of the TQA are:
Baulk Solenoids The baulk solenoids prevent the thrust levers from moving beyond the reverse idle position.
•
Thrust levers
•
ENG switches
•
RVDTs
Thrust Reverser Stow/Deploy Switches
•
Baulk solenoids
•
Thrust reverser stow/deploy switches
•
Autothrottle controller
The thrust reverser stow/deploy switches are single pole double throw type. They are activated when the thrust levers move into the reverse range.
•
Autothrottle servomotors
Autothrottle Controller The autothrottle controller provides output to the autothrottle servomotor.
Thrust Levers The thrust levers travel a total of 67° from the MAX REV to MAX positions. the thrust levers have hard stops at three positions; IDLE, MAX and MAX REV. A friction clutch provides a feel force and maintains the thrust lever in the selected position. The levers have an autothrottle disconnect takeoff/go-around (TOGA) switch built in.
switch
and
Autothrottle Servomotors Each thrust lever has an autothrottle servomotor that moves the thrust levers in response to the autothrottle controller commands.
a
Each thrust lever has a finger lift that allows the lever to move from idle to the reverse thrust position. ENG Switches The L(R) ENG switches initiate the engine start during an automatic start. During a manual start, they command fuel and ignition on. The L(R) ENG switches are selected OFF to shut down the engines.
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76-6
CS130_76_L2_EngineControl_ComponentInformation.fm
76 - Engine Controls 76-10 Engine Control
67° Autothrottle Disconnect Switches Finger Lifts
Takeoff/Go-Around Switches Baulk Solenoid
Thrust Reverser Stow/Deploy Switches
MAX IDLE
Servomotor REV
ENG Switches MAX REV Autothrottle Controller CS1_CS3_7600_003
Rotary Variable Displacement Transformers
Figure 3: Throttle Quadrant Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-7
CS130_76_L2_EngineControl_ComponentInformation.fm
76 - Engine Controls 76-10 Engine Control
DETAILED DESCRIPTION ENGINE CONTROL Each throttle lever mechanically drives three electrically independent RVDTs housed into one dual-coil and one single coil housing. Excitation voltage for the dual RVDTs is provided by the electronic engine control (EEC). Excitation voltage for the single RVDT is provided by the brake data concentrator unit (BDCU). The dual-coil RVDTs provide thrust lever position signal to the EEC. Channel A of the RVDT provides a signal for channel A of the EEC and channel B of the RVDT provides a signal for channel B of the EEC. The third RVDT (single coil housing) interfaces with the BDCU to provide electrically independent, digitized thrust lever angle information. This information is available to all systems, including the power plant, through the data concentrator unit module cabinet (DMC). A L (R) THROTTLE FAIL caution message is displayed when the EEC cannot determine the throttle position. The engine thrust is reduced to idle and the engine does not respond to thrust lever movement. THRUST REVERSER CONTROL The EEC provides a signal to control the thrust reverser (T/R) system baulk solenoid. When the T/Rs are deployed 85% of their travel, the EEC releases the baulk solenoid. This releases the thrust lever and allows full reverse thrust to be applied. The TQA baulk solenoids receive power from the aircraft control and distribution cabinets (CDCs). The EEC provides both the logic and the ground for the solenoids operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The interlock, provided by the baulk solenoids, incorporates an override operation in both directions that requires 25 lb of force per lever to overcome. The TQA has two stow/deploy microswitches per lever that control 28 VDC power to the thrust reverser track lock solenoids by providing the ground in circuit. These switches are hardwired to their respective control and distribution cabinets (CDCs). If either the left or right thrust lever is moved into the reverse thrust range while the aircraft is in the air, a THROTTLE IN REVERSE caution message is displayed. AUTOTHROTTLE CONTROL The autothrottle controller receives power from the flight control panel (FCP) when autothrottle is engaged. When engaged, the autothrottle software sends throttle control information to the autothrottle controller inside the TQA. The autothrottle controller then sends the commands to the autothrottle servomotors inside the TQA to move the thrust levers. The thrust levers are equipped with an autothrottle disconnect switch to disengage the autothrottle. THRUST LEVER WARNING If the thrust lever is not at idle when the engine is being started, the EEC initiates an aural warning to warn the crew, but does not prevent the engine from starting. The aural warning sounds “THRUST LEVER” three times.
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76-8
CS130_76_L3_EngineControl_DetailedDescription.fm
76 - Engine Controls 76-10 Engine Control
TQA
DMC 1 and DMC 2
LEGEND BDCU CDC DMC EEC FCP IFPC RIU RVDT TQA T/R
Autothrottle Disconnect Switch
Autothrottle
FCP Autothrottle Controller
28 VDC
Autothrottle Servomotor Position Signal RIU 1 and RIU 2
BDCU 1
Excitation Voltage
Single Channel RVDT DUAL-CHANNEL RVDT Position Signal Channel A
Brake Data Concentrator Unit Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Electronic Engine Control Flight Control Panel Integrated Fuel Pump and Control Radio Interface Unit Rotary Variable Differential Transformer Throttle Quadrant Assembly Thrust Reverser ARINC 429 Discrete
EEC Channel A
Excitation Voltage
IFPC
Position Signal Channel B CDC 3
Channel B
Excitation Voltage
BAULK SOLENOID
TQA L REV LOCK DC SSPC 3A BUS 1
Ground
Logic: Always On “THRUST LEVER, THRUST LEVER, THRUST LEVER”
L Engine Run Switch T/R Stow/ Deploy Switches
28 VDC
NOTE Left engine throttle operation shown. Right engine throttle operation similar.
To Track Lock Unit Solenoids
CS1_CS3_7600_011
CDC 1
Figure 4: Engine Control System Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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76-9
CS130_76_L3_EngineControl_DetailedDescription.fm
76 - Engine Controls 76-10 Engine Control
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages associated with the engine control system.
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TECHNICAL TRAINING MANUAL
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76-10
CS130_76_L2_EngineControl_MonitoringTests.fm
76 - Engine Controls 76-10 Engine Control
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L THROTTLE FAIL
Left throttle position is not recognized by the EEC.
R THROTTLE FAIL
Right throttle position is not recognized by the EEC.
THROTTLE IN REVERSE
Left or right thrust reverser selected in flight.
Table 2: ADVISORY Messages MESSAGE
LOGIC
L ENGINE FAULT Loss of redundant or non-critical function for the left engine. R ENGINE FAULT Loss of redundant or non-critical function for the right engine.
Table 3: STATUS Messages MESSAGE
LOGIC
L ENG SHUTDOWN
In-flight pilot commanded engine shut-down. Left engine run switch to OFF).
R ENG SHUTDOWN
In-flight pilot commanded engine shut-down. Right engine run switch to OFF.
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76-11
CS130_76_L2_EngineControl_MonitoringTests.fm
76 - Engine Controls 76-10 Engine Control
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TECHNICAL TRAINING MANUAL
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76-12
CS130_76_ILB.fm
ATA 77 - Engine Indicating
BD-500-1A10 BD-500-1A11
77 - Engine Indicating
Table of Contents 77-11 Engine Power Indicating System..................................77-2 General Description ..........................................................77-2 Component Location ........................................................77-4 Nf Speed Sensor..........................................................77-4 N1 Speed Probe ..........................................................77-4 Permanent Magnet Alternator Generator.....................77-4 Component Information ....................................................77-6 N1 Speed Probe ..........................................................77-6 Permanent Magnet Alternator Generator.....................77-8 Controls and Indications .................................................77-10 N1 Indication ..............................................................77-10 N1 Thrust Reference Bug ..........................................77-12 N1 Thrust Reference..................................................77-12 N1 Thrust Command Cursor......................................77-12 Takeoff Reference without FLEX or MAX Climb........77-12 N2 Indication ..............................................................77-14 Detailed Description .......................................................77-16 Monitoring and Tests .....................................................77-18 CAS Messages ..........................................................77-19
Monitoring and Tests .................................................... 77-28 CAS Messages ......................................................... 77-29 77-32 Vibration Monitoring................................................... 77-30 General Description ....................................................... 77-30 Component Location ..................................................... 77-32 Forward Vibration Sensor ......................................... 77-32 Aft Vibration Sensor .................................................. 77-32 Prognostics Health and Management Unit ............... 77-32 Data Storage Unit ..................................................... 77-32 Controls and Indications ................................................ 77-34 Indications ................................................................. 77-34 Detailed Description ...................................................... 77-36 Monitoring and Tests .................................................... 77-38 CAS Messages ......................................................... 77-39 Practical Aspects ............................................................ 77-40 Onboard Maintenance System.................................. 77-40 Fan Trim Balance Using the PW 1500G Trim Balance Helper ................................................. 77-58
77-21 Temperature Indicating System..................................77-20 General Description ........................................................77-20 Component Location ......................................................77-22 Exhaust Gas Temperature Probe and Cable Assembly .........................................................77-22 Oil Temperature Probe ..............................................77-22 Probe Junction ...........................................................77-22 Controls and Indications .................................................77-24 Indications..................................................................77-24 Detailed Description .......................................................77-26 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-i
CS130_M5_77_EngineIndicatingTOC.fm
77 - Engine Indicating
List of Figures Figure 1: Engine Power Indicating System...........................77-3 Figure 2: Engine Power Indicating System Component Location .............................................77-5 Figure 3: N1 Speed Probe....................................................77-7 Figure 4: Permanent Magnet Alternator Generator .............77-9 Figure 5: N1 Indication .......................................................77-11 Figure 6: N1 Thrust Reference Bug, N1 Thrust Reference, N1 Thrust Command Cursor, and Takeoff Reference........................................77-13 Figure 7: N2 Indication .......................................................77-15 Figure 8: Engine Power Indicating System Detailed Description ............................................77-17 Figure 9: Exhaust Gas Temperature Indicating System ................................................77-21 Figure 10: Exhaust Gas Temperature Indicating System Component Location..............77-23 Figure 11: Exhaust Gas Temperature Indications ................77-25 Figure 12: Exhaust Gas Temperature Detailed Description ............................................77-27 Figure 13: Vibration Monitoring ............................................77-31 Figure 14: Vibration Monitoring Component Location .........77-33 Figure 15: Vibration Monitoring Indications ..........................77-35 Figure 16: Vibration Monitoring Detailed Description ...........77-37 Figure 17: Trim Balance .......................................................77-41 Figure 18: Fan Hub Balance Weights...................................77-43 Figure 19: Trim Balance Procedure Main Menu...................77-45 Figure 20: Review VIB Data History .....................................77-47 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 21: Figure 22: Figure 23: Figure 24: Figure 25: Figure 26: Figure 27: Figure 28: Figure 29: Figure 30: Figure 31:
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Review/Edit Stored Weights ............................... 77-49 Perform Balance ................................................. 77-51 Display Balance Coefficients .............................. 77-53 Engine Serial Number Coefficients..................... 77-55 Modify Settings Screen....................................... 77-57 Fan Trim Balance ............................................... 77-59 Onboard Maintenance System Data................... 77-61 Monitor Engine Vibration .................................... 77-63 Fan Hub Balance Weights .................................. 77-65 Review Vibration Data ....................................... 77-67 PW1500G Trim Balance Helper Tool ................. 77-69
77-ii
CS130_M5_77_EngineIndicatingLOF.fm
77 - Engine Indicating
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77-iii
CS130_M5_77_EngineIndicatingLOF.fm
ENGINE INDICATING - CHAPTER BREAKDOWN ENGINE INDICATING CHAPTER BREAKDOWN
Speed (Power)
1
Temperature Indications
2
Vibration Monitoring
3
77 - Engine Indicating 77-11 Engine Power Indicating System
77-11 ENGINE POWER INDICATING SYSTEM GENERAL DESCRIPTION The engine power indicating system uses engine speed as a reference for engine power. The engine has three speed sensors that monitor engine speed: •
Fan speed probe
•
N1 speed probe
•
N2 speed sensor
The fan speed probe reads engine fan speed from the fan reduction gearbox front shaft gear. The N1 speed probe measures low-pressure compressor (LPC) speed. The N1 gauge has analog and digital readouts. N1 is the primary thrust control parameter. The thrust reference, calculated as an N1 value, is displayed by a reference bug on the analog display, and in digital format above the digital readout box. The N2 speed sensor measures high-pressure compressor (HPC) speed. The N2 speed sensor is part of the permanent magnet alternator generator (PMAG). The N2 is displayed in digital readout format only. The electronic engine control (EEC) receives the speed sensor signals. The EEC uses the signals for engine control and outputs indication signals for the N1 and N2 engine indication and crew alerting system (EICAS) display. The N1 gauges and the N2 digital readouts for both engines are displayed on EICAS. The fan speed is not displayed, but is used as backup in case of N1 signal failure.
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77-2
CS130_77_L2_EPIS_General.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
2 N1 SPEED PROBE
EEC
1 FAN SPEED PROBE
CLB 1
2 3
92.2
92.2
73.0
73.3
N1
718
722
EGT
86.5 1090 124 116
LEGEND
EEC PMAG
N1 (LPT and LPC) N2 (HPT and HPC) Nf (fan rotor) Electronic Engine Control Permanent Magnet Alternator Generator Analog ARINC 429
N2 FF (KPH) OIL TEMP OIL PRESS EICAS
86.6 1110 124 116
CS1_CS3_7700_005
3 PMAG
Figure 1: Engine Power Indicating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-3
CS130_77_L2_EPIS_General.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
COMPONENT LOCATION The engine power indicating system consists of the following components: •
Fan speed probe
•
N1 speed probe
•
Permanent magnet alternator generator (PMAG)
NF SPEED SENSOR The fan speed probe is located at the 1 o’clock position in the no.1 bearing support housing of the engine fan. N1 SPEED PROBE The N1 speed probe is mounted at station 2.5, at the rear of the compressor intermediate case at the 4:30 position. PERMANENT MAGNET ALTERNATOR GENERATOR The N2 speed coil is integral to the PMAG, which is located on the aft side of the main gearbox at the 7 o’clock position.
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77-4
CS130_77_L2_EPIS_ComponentLocation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
A
B
C
Fan Speed Probe
Permanent Magnet Alternator Generator
B N1 SPEED PROBE
C MAIN GEARBOX
CS1 CS3 7711 001 CS1_CS3_7711_001
A TOP VIEW
Figure 2: Engine Power Indicating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-5
CS130_77_L2_EPIS_ComponentLocation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
COMPONENT INFORMATION N1 SPEED PROBE The N1 speed probe is a dual sensor with two independent, isolated coils, and electrical connectors. The probe provides dual signals proportional to the rotational speed of the no. 2 bearing coupling nut located on the LPC. The EEC uses one signal per channel to generate the N1 value.
NOTE The N1 probe is approximately 20 cm (8 in.) in length. Care should be taken during removal and installation to avoid damaging the tip.
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77-6
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
B
Sensor A
A N1 SPEED PROBE
B COUPLING NUT
CS1_CS3_7711_003
Compressor Intermediate Case
Figure 3: N1 Speed Probe Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-7
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
PERMANENT MAGNET ALTERNATOR GENERATOR The PMAG stator has two speed coils that transmit two independent high rotor (N2) speed signals to the EEC. The PMAG stator also has dedicated coils for both EEC channel power generation, and fly-by-wire (FBW) power. The PMAG has offset captive bolts and transfer tubes to ensure proper clocking upon installation.
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77-8
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
Cover
Rotor
PERMANENT MAGNET ALTERNATOR GENERATOR STATOR
LEGEND EEC Channel A Power EEC Channel B Power Fly-By-Wire Power Speed (N2) Coils
CS1_CS3_7711_002
MAIN GEARBOX
Figure 4: Permanent Magnet Alternator Generator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-9
CS130_77_L2_EPIS_ComponentInformation.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
CONTROLS AND INDICATIONS N1 INDICATION The N1 indication consists of an analog and digital readout. A low-pressure compressor (LPC) speed of 100% N1 is equivalent to 10,600 rpm. The analog gauge consists of a pointer in a round arc, representing the operating range of the engine. The area below the pointer is shaded to provide a quick visual reference. At the upper end of the arc is a red line limit, indicating 100% N1 speed. The N1 digital display is in a white box to the right of the analog gauge. If the N1 speed is greater than 100%, the pointer moves above the red line limit. The pointer and the shaded area below changes to red. The arc extends above the red line and remains there even if the overspeed condition no longer exists. This indicates that the engine speed has exceeded 100%. The digital display and box color changes to red when N1 is above 100%. The display and box changes to white if the overspeed condition no longer exists. The N1 thrust reference bug readout is located on the outside of the analog N1 gauge arc.
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77-10
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
N1 REDLINE
Normal red line limit Overspeed Condition N1 overspeed condition
CLB N1 Red Line Limit
92.2
92.2
73.0
104
Takeoff Reference Without FLEX or MAX Climb
N1 N1 DIGITAL PARAMETER Symbol
EICAS
Description
86.5
Engine operation in normal range.
102
Engine operation above red line.
èèè
Invalid
CS1_CS3_7711_005
N1 Analog Pointer N1 Numeric Indication
Figure 5: N1 Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77-11
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
N1 THRUST REFERENCE BUG Other indications on the N1 display include an N1 thrust reference bug. The thrust reference bug is the best N1 for the selected thrust reference mode. The bug is displayed in magenta if the flight management system (FMS) calculates the thrust reference, or is displayed in cyan if it is entered manually into the FMS. N1 THRUST REFERENCE The N1 thrust reference is shown above the N1 digital display box. This value is displayed in magenta when calculated by the FMS, and in cyan when entered manually into the FMS. N1 THRUST COMMAND CURSOR The N1 thrust command cursor indicates the difference between the pointer and predicted N1 speed after the engine throttle is moved. When in automatic power reserve (APR) mode, the thrust reference bug changes to green. An offset remains in steady-state since the actual N1 is higher due to APR. Thrust Limitation at Low Speed The upper end of the thrust command cursor is shaded gray when thrust limitation at low speed is enabled during takeoff. The upper end of the gray area indicates the commanded N1, while the lower end indicates the thrust limitation. As the airspeed increases the gray area reduces and is removed when airspeed is above 80 kt. TAKEOFF REFERENCE WITHOUT FLEX OR MAX CLIMB The white radial line shows the thrust reference if FLEX was canceled. If FLEX was not used, the radial line is masked by the pointer. In climb, the white radial line shows the maximum climb thrust reference. If the information is invalid, the line is removed.
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77-12
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
Takeoff Reference if FLEX Was Canceled or MAX Climb Thrust Reference N1 Thrust Reference Bug
N1 THRUST REFERENCE BUG
N1 Thrust Reference
Symbol
Condition
Target N1 value entered manually in N1 data entry field
92.2
73.0
Bug position indicates target N1 value as computed automatically by the FMS
Thrust Command Cursor N1 SPEED N1 THRUST COMMAND CURSOR Symbol
Condition
Commanded decrease in N1
90.9
57.3 Commanded increase in N1 Thrust Limitation at Low Speed
APR APR Active
CS1_CS3_7711_006
COMMANDED INCREASE IN N1
Figure 6: N1 Thrust Reference Bug, N1 Thrust Reference, N1 Thrust Command Cursor, and Takeoff Reference Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-13
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
N2 INDICATION The N2 indication is a three-digit value located in the EICAS secondary engine indications area. The three-digit value represents a percentage of the N2 speed. In normal operation the digits are displayed in white. When N2 speed exceeds 100% the color changes to red.
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77-14
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
N2
86.6 1110 124 116
Symbol
Description
86.5
Engine operation in normal range
103.5
N2 above N2 red line threshold
èèè
Invalid
CS1_CS3_7711_007
EICAS
Figure 7: N2 Indication Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-15
CS130_77_L2_EPIS_ControlsIndications.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
DETAILED DESCRIPTION Each channel of the electronic engine control (EEC) receives the following inputs: •
Nf speed
•
N1 speed
•
N2 speed
As long as the EEC is receiving one of the N1 speed signals, the EEC uses N1 as the primary thrust reference. In the event that one EEC channel fails, the other channel maintains engine control.
Data concentrator unit module cabinet 1 (DMC 1) and DMC 2 also receive calculated N1 and N2 information via an ARINC 429 BUS for distribution to the aircraft systems. A L (R) ENG EXCEEDANCE warning message indicates overspeed of the N1 or N2 speeds. If the EEC detects a N1 or N2 speed in excess of 105%, the EEC initiates an automatic shutdown of the engine. In the event the N2 value drops below 5% of the engine governed speed, an L(R) ENG FAIL caution message is displayed.
In the event of a loss of the both N1 speed signals, the EEC calculates the N1 speed using the Nf signal. In the case of failure of one of the dual N1 speed signals, a L (R) ENGINE FAULT advisory message indicates a fault affecting the dispatch of the aircraft. The EEC uses the N2 speed signal for control of the fuel and ignition during starting and overspeed monitoring. In the event that one EEC channel fails, the other channel provides N2 speed. In the event of a loss of one of the N2 signals, a L(R) ENG OPER DEGRADED caution message is displayed on the engine indication and crew alerting system (EICAS). A L(R) ENG FAULT - FADEC FAULT 1 crew alerting system (CAS) advisory message is also displayed and rapid thrust movement is to be avoided. If both N2 signals are lost, the EEC commands an engine shutdown and the L(R) ENG FAIL caution CAS message is displayed. The EEC sends the N1, N2, and Nf information to the prognostics and health management unit (PHMU) via CAN DATA BUS. The PHMU uses the N1 and N2 speeds for engine vibration monitoring. The Nf is used by the PHMU for engine fan trim balancing. The EEC sends the calculated N1 and N2 information directly to the EICAS via an ARINC 429 BUS. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-16
CS130_77_L3_EPIS_DetailedDescription.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
CLB FAN SPEED PROBE Prognostics and Health Management Unit
Nf Speed
92.2
92.2
73.0
73.3
N1 EEC
N1 SPEED PROBE
N1 Speed 1 N1 Speed 2
718
Channel A
722
EGT PMAG
Channel B N2 Speed 1 N2 Speed 2
DMC 1 and DMC 2
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
LEGEND
EICAS
Data Concentrator Unit Module Cabinet Electronic Engine Control Permanent Magnet Alternator Generator Analog ARINC 429 CAN BUS
NOTE Left engine shown. Right engine similar.
CS1_CS3_7711_004
DMC EEC PMAG
86.6 1110 124 116
Figure 8: Engine Power Indicating System Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-17
CS130_77_L3_EPIS_DetailedDescription.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and fault messages for the engine power indicating system.
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77-18
CS130_77_L2_EPIS_MonitoringTests.fm
77 - Engine Indicating 77-11 Engine Power Indicating System
CAS MESSAGES
Table 3: FAULT Messages Table 1: WARNING Messages
MESSAGE DUAL ENG FAIL
MESSAGE
LOGIC L ENG and R ENG switches are ON and dual engine flameout.
LOGIC
L(R) ENG FAULT- INFO message will be set when any non-dispatchable FADEC FAULT 1 fault has been detected by the left (right) engine FADEC system or the DMC receives no valid data from left (right) engine for 10 seconds.
Table 2: CAUTION Messages MESSAGE
LOGIC
L ENG EXCEEDANCE
L N1 or L N2 or L ITT or L oil temperature above threshold.
R ENG EXCEEDANCE
R N1 or R N2 or R ITT or R oil temperature above threshold.
L ENG FAIL
L ENG switch in ON with the engine in a sub-idle condition and a relight or restart was unsuccessful.
R ENG FAIL
R ENG switch ON with the engine in a sub-idle condition and a relight or restart was unsuccessful.
L ENG OPER DEGRADED
Left EEC has detected a loss of thrust or degraded thrust control due to:
R ENG OPER DEGRADED
•
Fuel flow or actuator not tracking to commanded position
•
Input parameter in failsafe
•
Engine flameout
Right EEC has detected a loss of thrust or degraded thrust control due to: •
Fuel flow or actuator not tracking to commanded position
•
Input parameter in failsafe
•
Engine flameout
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77-19
CS130_77_L2_EPIS_MonitoringTests.fm
77 - Engine Indicating 77-21 Temperature Indicating System
77-21 TEMPERATURE INDICATING SYSTEM GENERAL DESCRIPTION The temperature indicating system measures the exhaust gas temperature (EGT) for display on the engine indication and crew alerting system (EICAS). EGT is monitored at station 5 of the engine. Two EGT probe and cable assemblies are used to transmit the EGT signals to the main oil temperature (MOT) probe and then to the electronic engine control (EEC). The oil temperature probe is used as the cold junction for the EGT signal. The EEC then calculates the EGT value and sends the information to the EICAS display. Each EGT probe and cable assembly consists of four EGT probes. The assemblies are semi-rigid and can only be replaced as an assembly. Individual EGT probes are not replaceable.
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77-20
CS130_77_L2_TIS_General.fm
77 - Engine Indicating 77-21 Temperature Indicating System
CLB Left EGT Probe and Cable Assembly
92.2
92.2
73.0
73.3
N1
OIL TEMPERATURE PROBE
718
ELECTRONIC ENGINE CONTROL
EGT
86.5 1090 124 116
Right EGT Probe and Cable Assembly
722
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
CS1_CS3_7721_001
EICAS
LEGEND EGT
Exhaust Gas Temperature ARINC 429
Figure 9: Exhaust Gas Temperature Indicating System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-21
CS130_77_L2_TIS_General.fm
77 - Engine Indicating 77-21 Temperature Indicating System
COMPONENT LOCATION The temperature indicating system consists of the following components: •
Exhaust gas temperature (EGT) probe and cable assembly
•
Oil temperature probe
•
Probe junction
EXHAUST GAS TEMPERATURE PROBE AND CABLE ASSEMBLY The EGT probe and cable assemblies are mounted around the circumference of the turbine exhaust case (TEC). There is one assembly on the left side of the engine, and one on the right side. OIL TEMPERATURE PROBE The oil temperature probe is installed on the oil control module (OCM) located on the right side of the engine. PROBE JUNCTION The probe junction is located on each side of the TEC at the 3 o’clock and 9 o’clock position.
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77 - Engine Indicating 77-21 Temperature Indicating System
A
B
EGT Probe and Cable Assembly
B PROBE JUNCTION
OIL TEMPERATURE PROBE
CS1_CS3_7721_006
A
Figure 10: Exhaust Gas Temperature Indicating System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77-23
CS130_77_L2_TIS_ComponentLocation.fm
77 - Engine Indicating 77-21 Temperature Indicating System
CONTROLS AND INDICATIONS INDICATIONS The exhaust gas temperature (EGT) display shows the EGT for each engine on the EICAS display, as calculated by the EEC. The gauges consist of an analogue and a four-digit display. The display is shown in degrees Celsius (°C). The analogue gauge consists of a pointer in a round arc representing the operating range of the engine. The area below the pointer is shaded to provide a quick visual reference. At the upper end of the arc is a red line limit, indicating maximum operating temperature. There is also a yellow line to indicate approaching maximum temperature limit. The electronic engine control (EEC) calculates an engine operating time limit in the yellow range. If the time limit exceeds the EGT needle, the shaded area and digital readout change to red. The yellow line is removed during takeoff, go-around, reverse, or when invalid. When the EGT pointer goes above the yellow line, the pointer and the shaded area below the pointer change to yellow. When the EGT pointer goes above the red line limit, the pointer and the shaded area below the pointer change to red. The arc is also extended in red above the red line limit, and stays there even if the overtemperature condition goes away. This indicates that the engine temperature has exceeded its limit. The digital EGT display changes color to match the pointer and shaded area.
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77 - Engine Indicating 77-21 Temperature Indicating System
EGT INDICATIONS
860 718
Engine start on ground only Yellow limit removed after engine start completed
722
EGT
1036
EICAS
EGT above amber line Engine start on ground only
Normal Indications - Engine Running on ground after start and in flight
1064
CS1_CS3_7721_005
EGT above red line
Figure 11: Exhaust Gas Temperature Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-25
CS130_77_L2_TIS_ControlsIndications.fm
77 - Engine Indicating 77-21 Temperature Indicating System
DETAILED DESCRIPTION The four signals from both EGT probe and cable assemblies are electrically averaged into two signals, which are sent to the oil temperature probe for compensation. The oil temperature probe sends compensated EGT signals to the EEC. The left EGT probe and cable assembly signal is sent to channel A. The signal of the right EGT probe and cable assembly is sent to channel B of the EEC. The EEC converts the average EGT analog signal to a digital signal. The two signals are averaged into a single output, which is sent directly to the display unit (DU) for display on the EICAS. A L(R) ENGINE FAULT advisory message is displayed when one of the EGT probe and cable assemblies fails.
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77 - Engine Indicating 77-21 Temperature Indicating System
Left EGT Probe and Cable Assembly
OIL TEMPERATURE PROBE
ELECTRONIC ENGINE CONTROL
Probe Junction Channel A
Channel A
Channel B
Channel B
718
Probe Junction
EGT EICAS
Right EGT Probe and Cable Assembly
EGT
CS1_CS3_7721_004
LEGEND Exhaust Gas Temperature ARINC 429 CAN BUS
Figure 12: Exhaust Gas Temperature Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-27
CS130_77_L3_TIS_DetailedDescription.fm
77 - Engine Indicating 77-21 Temperature Indicating System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) messages for the engine temperature indicating system.
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77 - Engine Indicating 77-21 Temperature Indicating System
CAS MESSAGES Table 4: CAUTION Messages MESSAGE
LOGIC
L ENG EXCEEDANCE
L N1 or L N2 or L ITT or L oil temperature above threshold.
R ENG EXCEEDANCE
R N1 or R N2 or R ITT or R oil temperature above threshold.
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77 - Engine Indicating 77-32 Vibration Monitoring
77-32 VIBRATION MONITORING GENERAL DESCRIPTION The engine vibration indicating system monitors, records, and analyzes engine vibration before displaying this information on the engine indication and crew alerting system (EICAS). Engine speed sensors are used, in conjunction with vibration monitors, to help determine which rotor may be out of balance. The prognostics and health management unit (PHMU) compares engine speed data from the electronic engine control (EEC) with vibration data sent from the forward and aft rotor vibration sensors. The PHMU receives power from DC BUS 1. Data stored in the EECs data storage unit (DSU) over a period of flights can be used by the onboard maintenance system (OMS) to conduct engine vibration surveys. Vibration sensors are used on both N1 and N2 rotors to send signals that are proportional to engine vibration. The PHMU converts the signals from charge to voltage, filters them, and correlates the frequencies with the specific rotor. The N1 and N2 vibration signals are then sent to the EEC for display on EICAS. These signals are also used by the aircraft OMS to calculate solutions for correcting a fan imbalance.
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77 - Engine Indicating 77-32 Vibration Monitoring
CLB DC BUS 1
92.2
92.2
73.0
73.3
N1
DSU
718
722
EGT
Speed Sensing FWD VIBRATION SENSOR
86.5 1090 124 116
N2 FF (KPH) OIL TEMP OIL PRESS
86.6 1110 124 116
EICAS PROGNOSTICS AND HEALTH MANAGEMENT UNIT
ELECTRONIC ENGINE CONTROL
MAINT MENU EEC1-A-TRIM BALANCE PROCEDURE MAIN MENU 4/90
AFT VIBRATION SENSOR REVIEW VIB DATA HISTORY
REVIEW/EDIT STORED WEIGHTS
PERFORM BALANCE
DISPLAY BALANCE COEFFICIENTS
LEGEND DSU
MODIFY SETTINGS
Data Storage Unit Analog ARINC 429
Write Maintenance Reports
OMS
Cancel
CS1_CS3_7732_004
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
Figure 13: Vibration Monitoring Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_77_L2_VM_General.fm
77 - Engine Indicating 77-32 Vibration Monitoring
COMPONENT LOCATION The vibration monitoring system consists of the following components: •
Forward vibration sensor
•
Aft vibration sensor
•
Prognostics health and management unit (PHMU)
•
Data storage unit (DSU)
FORWARD VIBRATION SENSOR The forward vibration sensor is mounted on the compressor intermediate case at the 9 o’clock position. AFT VIBRATION SENSOR The aft vibration sensor is mounted on the low-pressure turbine (LPT) housing at the 3 o’clock position. PROGNOSTICS HEALTH AND MANAGEMENT UNIT The prognostics health and management unit (PHMU) is mounted to the fan case adjacent to the EEC. DATA STORAGE UNIT The data storage unit (DSU) is located on the EEC.
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77 - Engine Indicating 77-32 Vibration Monitoring
A
A AFT VIBRATION SENSOR RIGHT SIDE VIEW Electronic Engine Control
Data Storage Unit
Prognostics and Health Management Unit
LEFT SIDE VIEW
B
FWD VIBRATION SENSOR
CS1_CS3_7731_001
B
Figure 14: Vibration Monitoring Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-33
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77 - Engine Indicating 77-32 Vibration Monitoring
CONTROLS AND INDICATIONS INDICATIONS The following vibration indications are displayed in the EICAS: •
N1 vibration is displayed as an AMBER VIB flag in the associated N1 gauge
•
N2 vibration is displayed as an AMBER VIB flag adjacent to the digital readout
•
FAN VIB is displayed in units
The N1 VIB is displayed as a flag in the N1 gauge when the parameter exceeds the normal vibration threshold of 0.52 inches per second (IPS). The REV icon replaces the N1 VIB display when the thrust reversers deploy. The N2 VIB is displayed as a flag adjacent to the N2 speed when the parameter exceeds the normal vibration threshold of 1.2 IPS. The FAN VIB level is posted on EICAS in units from 0 to 8. The units represent a percentage of the vibration threshold where 4.0 is 100%. The normal vibration threshold for the engine fan is 1.2 IPS. FAN VIB is only displayed when an engine fan vibration exceeds 4.0 units, or a fault has occurred in the vibration monitoring system. When this occurs, both left and right engine FAN VIB is displayed. The fan vibration exceedance is displayed in amber, while normal indications are displayed in white.
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77 - Engine Indicating 77-32 Vibration Monitoring
CLB 92.2
92.2
73.0
73.3
FAN VIBRATION Symbol
N1
718
VIB
722
0.6 4.2 èèè
Description Vibration at or below high vibration threshold Vibration above high vibration threshold If both vibration monitors fail on one side.
EGT N1 and N2 VIBRATION
86.5 VIB N2 86.6 1090 FF (KPH) 1110 124 OIL TEMP 124 116 OIL PRESS 116 FAN VIB 4.2 0.4
Symbol
VIB
Description Vibration above high vibration threshold N1 VIB is not displayed when the thrust reverser is operating.
CS1_CS3_7732_001
EICAS
Figure 15: Vibration Monitoring Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
DETAILED DESCRIPTION Vibration monitoring is performed through data input from the two engine vibration sensors, as well as speed inputs from the EEC for N1, N2, and the fan speed (Nf). The PHMU receives engine speed signals from the EEC and vibration signals from the vibration sensors. The Nf speed signal is used in conjunction with the forward vibration sensor to measure fan vibration. The PHMU compares the engine speeds and vibration signals to determine the vibration level for the fan, N1, and N2 shafts. The PHMU sends the vibration signals to the EEC via a CAN BUS. The EEC sends vibration information for display on EICAS. When engine vibration exceedances are detected, an ENG VIBRATION caution message is displayed, indicating excessive vibration levels. Engine vibration information is also sent to the DMCs for the OMS and health management unit (HMU).
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77 - Engine Indicating 77-32 Vibration Monitoring
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
CDC 1-10-8 2 PHMU DC BUS 1
SSPC 5A
Logic: Always on
ELECTRONIC ENGINE CONTROL N1 Signal N2 Signal
Channel A
Channel B
EICAS
HMU
DMC 1 and DMC 2
OMS
Nf Signal
FWD VIBRATION SENSOR
AFT VIBRATION SENSOR
DMC
Data Concentrator Unit Module Cabinet ARINC 429 CAN BUS NOTE Left engine shown.
CS1_CS3_7731_002
LEGEND
Figure 16: Vibration Monitoring Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the engine vibration monitoring system.
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77 - Engine Indicating 77-32 Vibration Monitoring
CAS MESSAGES Table 5: CAUTION Messages MESSAGE ENGINE VIBRATION
LOGIC Left or right engine vibration (either fan vibration, N1 or N2).
Table 6: ADVISORY Message MESSAGE
LOGIC
L (R) ENG FAULT Loss of redundant or non-critical function for the left (right) engine. Refer to INFO messages.
Table 7: INFO Messages MESSAGE
LOGIC
77 L (R) ENGINE FAULT - PHMU INOP
FADEC logic has detected that the PHMU is failed.
77 L (R) ENGINE FAULT - N1/FAN VIBRATION NON DEGRADED
INFO message will be set when a forward vibration sensor fault has been detected or N1 or NF speed input for vibration processing has failed.
77 L (R) ENGINE FAULT - N2 VIBRATION NON DEGRADED
INFO message will be set when a rear vibration sensor fault has been detected or N2 speed input for vibration processing has failed.
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77 - Engine Indicating 77-32 Vibration Monitoring
PRACTICAL ASPECTS ONBOARD MAINTENANCE SYSTEM Trim Balance The purpose of the trim balance procedure is to provide the information needed to balance the fan and/or low spool of the engine. The fan is a large rotating mass, even minor blade damage can cause an imbalance. A trim balance procedure is used to reduce fan-related vibration onwing. While the engine is running, the prognostics and health management unit (PHMU) continuously collects data for trim balancing using values from the vibration sensors and the predetermined Nf, N1, and N2 engine speeds. Historical vibration data stored in the EECs data storage unit (DSU) is used as reference in calculating the trim balance solutions. The system corrects small amounts of residual imbalance, which could be related to minor blade damage or fan blade replacement, even when matched sets are replaced.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU EEC1-A-TRIM BALANCE PROCEDURE MAIN MENU 4/90
REVIEW VIB DATA HISTORY Data Storage Unit
REVIEW/EDIT STORED WEIGHTS DMC
PERFORM BALANCE
DISPLAY BALANCE COEFFICIENTS ELECTRONIC ENGINE CONTROL
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT PROGNOSTICS AND HEALTH MANAGEMENT UNIT
LEGEND DMC
Write Maintenance Reports
Data Concentrator Unit Module Cabinet
ONBOARD MAINTENANCE SYSTEM
Cancel
CS1_CS3_7730_001
MODIFY SETTINGS
Figure 17: Trim Balance Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-41
CS130_77_L2_VM_PracticalAspects.fm
77 - Engine Indicating 77-32 Vibration Monitoring
Fan Hub Balance Weights N1 (fan rotor) vibration signals are sent to the PHMU and the EEC for interpretation and calculation of optimal trim balance solutions. The onboard maintenance system (OMS) uses these solutions to indicate where to install balance weights on the fan rotor disc. Each weight has a different part number, based on its specific weight. The fan speed sensor is used to determine the fan index position sensing. It measures the fan index rotational position, once-per-revolution, to allow the location of vibration imbalances to support the fan trim balance. The trim balance procedure involves running a baseline (as-is) vibration survey. The final correction weight is then installed on a balance weight flange, located on the front of the fan rotor hub. The software-stored weight configuration should exactly match the actual balance weight configuration on the engine in order for the OMS system-provided balance solution to be correct. Reviewing and editing the stored balance weight configuration is a precondition for performing a trim balance. A 0° reference point is aligned with a dimple on the fan hub shaft. Trim balance weights are installed on either side of the module balance weights. Module balance weights are riveted in place during production or engine rebuild, and should not be disturbed during trim balance procedures.
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77 - Engine Indicating 77-32 Vibration Monitoring
0° Reference for Weight Angles Aligned With a Shaft Dimple
Fan Drive Shaft Dimple Location
Trim Balance Locations (for clip-on weights)
VIEW FORWARD LOOKING AFT (fan drive shaft shown)
CS1_CS3_7730_009
Z-Balance and Module Balance Locations (for riveted weights)
Figure 18: Fan Hub Balance Weights Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Trim Balance Procedure Main Menu On the OMS maintenance page, the trim balance procedure main menu screen provides an entry point into the trim balance subprocedures, including: •
REVIEW VIB DATA HISTORY
•
REVIEW/EDIT STORED WEIGHTS
•
PERFORM BALANCE
•
DISPLAY BALANCE COEFFICIENTS
•
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
•
MODIFY SETTINGS
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU EEC1-A-TRIM BALANCE PROCEDURE MAIN MENU 4/90
REVIEW VIB DATA HISTORY
REVIEW/EDIT STORED WEIGHTS BEFORE A BALANCE SOLUTION IS IMPLEMENTED THE CURRENT WEIGHT CONFIGURATION STATUS SHOULD BE REVIEWED AND CORRECTED IF NEEDED.
PERFORM BALANCE BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM LAST COMPLETED RUN OR BASED ON SEVERAL PAST RUNS.
DISPLAY BALANCE COEFFICIENTS
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
MODIFY SETTINGS
Write Maintenance Reports
Cancel
ONBOARD MAINTENANCE SYSTEM
CS1_CS3_7730_002
THIS PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY BE USED IF THE GENERIC COEFFICIENTS DO NOT PRODUCE SATISFACTORY RESULTS.
Figure 19: Trim Balance Procedure Main Menu Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Review VIB Data History REVIEW VIB DATA HISTORY determines changes to the overall vibration levels in the recent flight history.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
EEC1-A-VIB HISTORY MENU 4/90 0
10/90 -----------TRIM BALANCE RELATED VIB HISTORY-----------
REVIEW VIB DATA HISTORY
FAN VIB HISTORY
REVIEW/EDIT STORED WEIGHTS GHTS LPC VIB HISTORY
LPT VIB HISTORY
PERFORM BALA BALANCE BALANC ALAN NCE E
----------OTHER BALANCE RELATED VIB HISTORY----------
HP SPOOL VIB HISTORY DISPLAYY BALANCE COEFFICIENTS BROADBAND VIB HISTORY CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT WEIGHT ----VIEW A SUMMARY OF AVAILABLE BALANCE SOLUTIONS----
MODIFY SETTINGS GS
RETURN TO TRIM BALANCE MAIN MENU
Write Maintenance M Reports ports ts
Cance Cancel ncel l
Write Maintenance Reports
Cancel
CS1_CS3_7730_003
VIEW BALANCE SOLUTION SUMMARY
Figure 20: Review VIB Data History Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Review/Edit Stored Weights REVIEW/EDIT STORED WEIGHTS provides a starting point for examining and changing software stored balance weight configurations for each of the correction planes (fan, LPC, and LPT).
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
GPS 1-REVIEW/EDIT STORED WEIGHT CONFIGURATION 4/90 0 THIS PROCEDURE WILL REQUIRE ACCESS TO BALANCE CORRECTION PLANES TO EXAMINE CURRENT STATE.
REVIEW VIB DATA HISTOR HISTORY TORY Y
REVIEW/EDIT STORED WEIGHTS
REVIEW/EDIT FAN CORRECTION PLANE WEIGHTS
PERFORM BALA BALANCE BALANC ALAN NCE E
REVIEW/EDIT LPC CORRECTION PLANE WEIGHTS DISPLAYY BALANCE COEFFICIENTS REVIEW/EDIT LPT CORRECTION PLANE WEIGHTS
MODIFY SETTINGS GS
RETURN TO TRIM BALANCE MAIN MENU
Write Maintenance M Reports ports ts
Cance Cancel ncel l
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7730_004
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT WEIGHT
Figure 21: Review/Edit Stored Weights Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-49
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77 - Engine Indicating 77-32 Vibration Monitoring
Perform Balance PERFORM BALANCE provides options for selecting which type of balance procedure to perform. Each of the provided buttons lead to separate balance solutions, based on the selected balance plane (fan, LPC or LPT). All balance solutions are precalculated and stored during previous engine runs.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
EEC1-A-SELECT BALANCE PROCEDURE 4/90 0
4/90 THIS PROCEDURE WILL REQUIRE ACCESS TO BALANCE DO PLANES FAN TRIM BALANCECURRENT (SINGLE PLANE) CORRECTION TO EXAMINE STATE.
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y
PROCEDURE WILL REQUIRE FAN CORRECTION PLANE ACCESS.
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
DO LPT TRIM BALANCE (SINGLE PLANE)
PERFORM BALANCE OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM RO LAST ASSTT
PROCEDURE WILL REQUIRE LPT CORRECTION PLANE ACCESS.
DISPLAY BALANCE COEFFICIENTS IENTS
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT CA
MODIFY DIFY SETTINGS SETTINGS
Write Maintenance M Reports Repor ports ts
RETURN TO TRIM BALANCE MAIN MENU
Cance Cancel ancel l
Write Maintenance Reports
Cancel
CS1_CS3_7730_005
S PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY THIS S ICC COEFFICIENTS RODUCE BE USED PRODUCE IF THE GENERIC CCO OEFFI OOEFFICIENTS IC CI CI NNTSS DO DO NO NOT OT PRO PRODUC PPRODU ROD O UC CE SATISFACTORY RESULTS. .
Figure 22: Perform Balance Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77-51
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77 - Engine Indicating 77-32 Vibration Monitoring
Display Balance Coefficients The DISPLAY BALANCE COEFFICIENTS page is an advanced screen used to modify OMS trim balance settings. Refer to the Aircraft Maintenance Publication (AMP) for guidance as to when this advanced subprocedure should be used.
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
GPS 1-REVIEW/EDIT STORED WEIGHT CONFIGURATION 4/90 0
8/90
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y FAN GENERIC NUM MAG PHASE 1: X.X XXX 2: X.X XXX 3: X.X XXX 4: X.X XXX 5: X.X XXX 6: X.X XXX 7: X.X XXX 8: X.X XXX 9: X.X XXX 10: X.X XXX
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
PERFORM BALANC BALANCE BALAN NCE E
NUM 11: 12: 13: 14: 15: 16: 17: 18: 19: 20:
MAG X.X X.X X.X X.X X.X X.X X.X X.X X.X X.X
PHASE XXX XXX XXX XXX XXX XXX XXX XXX XXX XXX
OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM ROM RO LAST ASSTT
EACH SET CONTAIN 20 PAIRS OF COEFFICIENTS EACH PAIR = A SENSITIVITY FACTOR AND A LAG ANGLE
DISPLAY BALANCE COEFFICIENTS
E-FLANGE 1...................10
EACH BLOCK OF 10 PAIRS ARE PER THE 10 SPEED POINTS
S PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY THIS S ICC COEFFICIENTS RODUCE BE USED IF THE GENERIC PRODUCE CCO OEFFI OOEFFICIENTS IC CI CI NNTSS DO DO NO NOT OT PRO PRODUC PPRODU ROD O UC CE SATISFACTORY RESULTS. .
RETURN TO SELECT COEFFICIENT SCREEN
MODIFY SETTINGS SETTING ETTINGS S
Write ite Maintenance M Reports ports orts
Cance Cancel ncel l
Write Maintenance Reports
Cancel
CS1_CS3_7730_006
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT CA
P-FLANGE 11..................20
Figure 23: Display Balance Coefficients Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
Engine Serial Number Coefficients The calculated ESN coefficients, using the trial weight screen, should only be used if the normal procedure does not produce the desired reduction in vibration. The screen enables the installation of trial weights to help determine a solution.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
GPS 1-CALCULATE BALANCE COEFFICIENTS-1 4/90 0
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y
TO CALCULATE BALANCE COEFFICIENTS, THE OPERATOR MUST COMPLETE THE FOLLOWING STEPS IN ORDER 1) SELECT DESIRED BALANCE PLANE 2) REVIEW AND CORRECT CURRENT WEIGHT CONFIGURATION
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
3) ADD A TRIAL WEIGHT TO BALANCE PLANE 4) EXIT INTERACTIVE MODE TO CONDUCT A TRIM BALANCE SURVEY PER AMM
PERFORM BALANC BALANCE BALAN NCE E
5) RETURN TO TRIM BALANCE INTERACTIVE PROCEDURE TO CONFIRM REMOVAL OF TRIAL WEIGHT AND TO REVIEW NEW BALANCE COEFFICIENT
OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM ROM RO LAST ASSTT
DISPLAY BALANCE COEFFICIENTS OEFFICIENTS
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT
SELECT "CONTINUE" TO START STEP 1...
MODIFY SETTINGS SETTING ETTINGS S
Write ite Maintenance M Reports ports orts
RETURN TO TRIM BALANCE MAIN MENU
Cance Cancel ncel l
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7730_007
THIS PROCEDURE RE REQUIRES ROUND RUN HOULD QUIRES A GRO GROUND RU AND SSHOULD OULD ONLY O S ICC COEFFICIENTS OEFFICIENTS S DO RODUCE BE USED PRODUCE IF THE GGENERIC CCO OOEFFI IC CI CI NNTS DO NO NOT OT PRO PRODUC PPRODU ROD O UC CE SATISFACTORY RESULTS. .
Figure 24: Engine Serial Number Coefficients Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Modify Settings Screen The MODIFY SETTINGS screen is an advanced screen that allows setting changes for trim balance processing, clearing configuration, and balance histories.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT T MENU MENU
MAINT MENU
EEC1-A-TRIM EEC1-A -AA-TRIM BBALANCE ALLANCE PRO PROCEDURE ROCEDURE MA MAIN MENUU
REVIEW EVIEW VIB DATA DATA HISTOR HIS HISTORY TORY Y
GPS 1-MODIFY TRIM BALANCE SETTINGS 1/90 * NOTE * TRIM BALANCE SETTING ARE USED TO CONFIGURE SOFTWARE PROCESSING WITH THE SELECTIONS PROVIDED BELOW.
REVIEW/EDIT EVIEW/EDIT STORED WEIGHTS
CLEAR BALANCE VECTOR HISTORY
4/90 0
BALANCE DATA FROM PAST RUNS WILL NOT BE USED FOR FUTURE BALANCE SOLUTION CALCULATIONS
PERFORM BALANC BALANCE BALAN NCE E
EXCLUDE GROUND RUN FROM UPDATING HISTRY
OM LLAS BALANCE CAN BE PERFORMED BASED ON VIB DATA FROM ROM RO LAST ASSTT
PREVENT NORMAL OPERATION GROUND RUNS FROM UPDATING BALANCE VECTOR HISTORY
DISPLAY BALANCE COEFFICIENTS OEFFICIENTS
CLEAR STORED VIB HISTORY
CALCULATE ESN COEFFICIENTS USING TRIAL WEIGHT CA
CLEAR ALL STORED BALANCE WEIGHT CONFIGURATION
MODIFY SETTINGS
RETURN TO TRIM BALANCE MAIN MENU
Write Maintenance M Reports ports orts
Cance Cancel ncel l
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7730_008
S PROCEDURE REQUIRES A GROUND RUN AND SHOULD ONLY THIS S ICC COEFFIC RODUCE COEFFICIENTS PRODUCE BE USED IF THE GENERIC CO COEFFICI O CIENTS IENTS NNTSS DO DO NOT NO PRO PPRODUC PRODU ROD O UC CE SATISFACTORY RESULTS. .
Figure 25: Modify Settings Screen Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
FAN TRIM BALANCE USING THE PW 1500G TRIM BALANCE HELPER Test Set Up The Pratt & Whitney Special Instruction no. 44F - 16 Rev. A provides instructions for performing an on-wing fan trim balance using the onboard maintenance system (OMS) and the PW1500G Fan Trim Balance Helper Tool. Use the OMS to get weight and angle data for the fan. Use this data with the PW1500G Trim Balance Helper Tool provided in this Special Instruction to develop a balance solution. Perform a baseline engine vibration survey as follows: 1. Start both engines and allow engines to idle for a minimum of 5 minutes to allow the engine to become thermally stable. Make sure all the engine parameters are within the allowable limits. 2. For the engine under test, slowly move the thrust lever forward to 67-69% N1 speed and let the engine become stable for 3 minutes. The engine not being tested can be operated as required during this survey. 3. For the tested engine, slowly move the thrust lever back to MINIMUM IDLE power and let the engine become stable for 3 minutes.
NOTE Do not run the tested engine at idle power for more than 3 minutes. The engine can cool and the thermal stabilization steps must be repeated.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77 - Engine Indicating 77-32 Vibration Monitoring
Data Storage Unit
DMC
ELECTRONIC ENGINE CONTROL
LEGEND DMC
Data Concentrator Unit Module Cabinet ONBOARD MAINTENANCE SYSTEM
CS1_CS3_7730_001
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
Figure 26: Fan Trim Balance Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Onboard Maintenance System Data Prior to performing a baseline vibration survey, access the onboard maintenance system (OMS) system as follows: 1. Select Perform LRU/Systems Operations 2. Select ATA 73 - Engine Fuel & Control 3. Select Left or Right Engine EEC for the engine under test 4. Select Vibration from the drop-down menu to monitor the engine vibration data At the completion of the vibration survey and prior to engine shutdown, Data Reader is selected to get the fan trim balance solution weight and solution angle.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU
MAINT MENU Maintenance Main Menu
View Post Flight Summary View Flight Deck Effects (5 Actives)
View System Exceedances (4
Stored)
LRU/System Operations
LRU/System Operations
LRU/System Operations
View
All ATAs
View
All ATAs
View
All ATAs
Select ATA
21-00-00 AIR CONDITIONING
Select ATA
73-00-00 ENGINE FUEL & CONTROL
Select ATA
73-00-00 ENGINE FUEL & CONTROL
Select LRU/System
Prev 38-00-00 44-00-00 46-00-00 47-00-00 49-00-00 52-00-00 73-00-00
Select LRU/System Left Engine EEC (1)
View Fault Messages (5 Actives) View Service Messages (0 Active)
MAINT MENU
DATA Data Reader
WATER/WASTE CABIN CONTROLLER INFORMATION SYSTEMS FUEL TANK INERTING APU-GENERAL DOORS ENGINE FUEL & CONTROL
DATA Actuators
Left Engine EEC (1) Left Engine EEC (1) channel B Right Engine EEC (2) Right Engine EEC (2) channel A Right Engine EEC (2) channel B
Select LRU/System Left Engine EEC (1)
DATA Actuators
Air Data
Air Data
Cowl Anti-Ice
Cowl Anti-Ice
Data Reader
Data Reader
Fuel System
Fuel System
Oil System
Oil System
Performance
Performance
Starting
Starting
Thrust Ratings
Thrust Ratings
Write Maintenance Reports
Thrust Reverser
Thrust Reverser
Utility Functions
Vibration
Vibration
View System Trends View Aircraft Life Cycle Data View System Parameters Perform LRU/System Operations
CS1_CS3_7730_010
View System Configuration Data
Figure 27: Onboard Maintenance System Data Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Monitor Engine Vibration The engine vibration data for N1, N2, and the fan Nf can be monitored during the engine run. For the engine being tested, slowly move the thrust lever to achieve the following settings: •
Perform a slow acceleration (40-45 sec) to 60 +/-1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 64 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 68 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 72 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 76 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 80 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow acceleration (5-10 sec) to 83 +/- 1% N1. Stabilize for 1 minute
•
Perform a slow deceleration (60-90 sec) to MINIMUM IDLE
NOTE The maximum speed may be limited by TLA rating. The ENG VIBRATION caution CAS can be disregarded for 30 seconds during this test in order to obtain vibration data as long as it is related to fan vibration.
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TECHNICAL TRAINING MANUAL
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT MENU LRU/System Operations View
All ATAs
Select ATA
73-00-00 ENGINE FUEL & CONTROL
Select LRU/System Left Engine EEC (1)
DATA Actuators Air Data Cowl Anti-Ice Data Reader Fuel System Oil System Performance Starting Thrust Ratings Thrust Reverser
CS1_CS3_7730_012
Vibration
Figure 28: Monitor Engine Vibration Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
77-63
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77 - Engine Indicating 77-32 Vibration Monitoring
Fan Hub Balance Weights A 0° reference point is aligned with a dimple on the fan hub shaft. Trim balance weights are installed on either side of the module balance weights. Module balance weights are riveted in place during production or engine rebuild, and should not be disturbed during trim balance procedures. The fan hub balance weights can be accessed as follows: 1. Remove the inlet cone cover from the inlet cone. 2. Rotate the reference dimple marked on the fan drive shaft to the top dead center (TDC) position (12:00). 3. Note and record any previously installed trim balance weights (spring loaded weights only) and hole location.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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77 - Engine Indicating 77-32 Vibration Monitoring
0° Reference for Weight Angles Aligned With a Shaft Dimple
Fan Drive Shaft Dimple Location
Trim Balance Locations (for clip-on weights)
VIEW FORWARD LOOKING AFT (fan drive shaft shown)
CS1_CS3_7730_009
Z-Balance and Module Balance Locations (for riveted weights)
Figure 29: Fan Hub Balance Weights Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
Review Vibration Data Access the Data Reader page to get the following data: •
Record the values displayed for the following parameters: -
Label 47 Fan Trim Balance Solution Weight
-
Label 54 Fan Trim Balance Solution Angle
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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77 - Engine Indicating 77-32 Vibration Monitoring
MAINT AINT MENU MENU
MAINT MENU
L EEC - Data Reader
Select ATA
All ATAs
Refre Refresh freshh
73-00-00 ENGINE FUEL & CONTROL
- LRU: Left Engine EEC (1) + Label: 43 N1 Vibration - Front Vib Senso + Label: 43 N1 Vibration - Front Vib Senso + Label: 44 N2 Vibration - Rear Vib Sensor + Label: 44 NF Vibration - Rear Vib Sensor + Label: 45 NF Vibration - Rear Vib Sensor + Label: 45 NF Vibration - Rear Vib Sensor + Label: 46 N1 Phase Angle - Front Vib Sen + Label: 47 N1 Phase Angle - Front Vib Sen - Label: 47 Fan Trim Balance Solution Weig SDI 1 Fan Trim Balance Solution Weight 0.0000 SSN [BNR] 3 - Label: 47 Fan Trim Balance Solution Weig SDI 1 Fan Trim Balance Solution Weight 0.0000 SSN [BNR] 3 - Label: 54 Fan Trim Balance Solution Angl SDI 1 Fan Trim Balance Solution Angle_ 0.0000 SSN [BNR] 3 - Label: 54 Fan Trim Balance Solution Angl SDI 3 Fan Trim Balance Solution Angle_ 0.0000 SSN [BNR] 3 + Label: 55 LPT Trim Balance Solution Weig + Label: 55 LPT Trim Balance Solution Weig + Label: 56 LPT Trim Balance Solution Angl + Label: 56 LPT Trim Balance Solution Angl + Label: 60 ODM Sensor Raw Signal RMS Valu
Select LRU/System Left Engine EEC (1)
DATA Actuators Air Data Cowl Anti-Ice Data Reader Fuel System Oil System Performance Starting Thrust Ratings Thrust Reverser Vibration
Write Mai Maintenance ce Reports orts
+ All
- AAll ll Bits Bits Bits Bits Bits Bits Bits Bits Bits
9-10 16-29 30-31 Bits
9-10 16-29 30-31 Bits
9-10 16-29 30-31 Bits
9-10 16-29 30-31 Bits Bits Bits Bits Bits
CS1_CS3_7730_011
LRU/System Operations View
RETURN TO RET LRU/SYS LR RU OPS PS S
Figure 30: Review Vibration Data Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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77 - Engine Indicating 77-32 Vibration Monitoring
PW1500G Trim Balance Helper Tool Using the tool instructions provided in the Excel off-line tool attached to the Special Instruction, calculate fan trim balance weight(s) and location information as follows: 1. Use Label 47 Fan Trim Balance Solution Weight_FANWGT and Label 54 Fan Trim Balance Solution Angle_FANANG values to determine the list of part numbers and placement based PW1500G Trim Balance Helper Tool. 2. Enter the previously installed trim balance spring loaded weights into the PW1500G Trim Balance Helper Tool. 3. Enter the Label 47 Fan Trim Balance Solution Weight_FANWGT value from the OMS in the Magnitude field of the PW1500G Trim Balance Helper Tool. 4. Enter the Label 54 Fan Trim Balance Solution Angle_FANANG value from the OMS in the Direction field of the PW1500G Trim Balance Helper Tool. 5. Click SET button. 6. Record results on the recommendation screen. 7. Remove or install fan trim balance weight(s) by size and location specified in the PW1500G Trim Balance Helper Tool. 8. Ensure weights are installed with spring forward. 9. Install the inlet cone cover to the inlet cone. Perform Test 8 - Vibration Analysis as per the Aircraft Maintenance Publication. Observe the engine vibration data during engine run with the OMS Further fan trim balance is not necessary once Nf vibration is under 0.5 IPS or 1.8 units. If the vibration survey data indicates that the fan vibration levels are unacceptable, repeat the procedure.
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TECHNICAL TRAINING MANUAL
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77 - Engine Indicating
CS1_CS3_7730_013
77-32 Vibration Monitoring
Figure 31: PW1500G Trim Balance Helper Tool Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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ATA 78 - Exhaust
BD-500-1A10 BD-500-1A11
78 - Exhaust
Table of Contents 78-36 Thrust Reverser Structure ............................................78-2 General Description ..........................................................78-2 Component Location ........................................................78-6 Inner Fixed Structure ...................................................78-6 Translating Sleeve .......................................................78-6 Component Information ....................................................78-8 Translating Sleeves .....................................................78-8 Blocker Doors ............................................................78-10 Cascades ...................................................................78-12 78-31 Thrust Reverser Actuation System.............................78-14 General Description ........................................................78-14 Component Location ......................................................78-16 Isolation Control Unit..................................................78-16 Directional Control Unit ..............................................78-16 Locking Feedback Actuators......................................78-18 Locking Actuators ......................................................78-18 Flexible Drive Shaft/Deploy Tubes.............................78-18 Manual Drive Units.....................................................78-18 Track Lock Valves......................................................78-18 Track Lock Units ........................................................78-18 Component Information ..................................................78-20 Isolation Control Unit..................................................78-20 Directional Control Unit ..............................................78-20 Locking Feedback Actuator ......................................78-22 Locking Actuator .......................................................78-24 Locking Feedback Actuator and Locking Actuator Installation..................................................................78-26 Track Lock Unit ..........................................................78-28 Track Lock Valve .......................................................78-30 Throttle Quadrant Assembly ......................................78-32 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Detailed Component Information ................................... 78-34 Locking Feedback Actuator and Locking Actuator ...................................................... 78-34 Throttle Quadrant Assembly Thrust Lever Control ................................................. 78-36 Controls and Indications ................................................ 78-38 Controls .................................................................... 78-38 Indications ................................................................ 78-40 Detailed Description ...................................................... 78-42 Thrust Reverser Locked............................................ 78-42 Isolation Control Unit Energized Actuators Overstow................................................... 78-44 Track Lock units Released........................................ 78-46 DCU Energized ......................................................... 78-48 Thrust Reversers Fully Deployed .............................. 78-50 Stow Command......................................................... 78-52 Thrust Reverser Stowed ........................................... 78-54 Thrust Reverser Actuating System Control Electrical Schematic ..................................... 78-56 Thrust Reverser Actuating System Control Logic ............................................................. 78-58 Monitoring and Tests ..................................................... 78-60 CAS Messages ......................................................... 78-61 Practical Aspects ............................................................ 78-62 Lockout the Thrust Reversers for Dispatch .............. 78-62 Manually Operate the Translating Sleeves for Maintenance ........................................................ 78-66 Thrust Reverser Cycling Test.................................... 78-68
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_M5_78_ExhaustTOC.fm
78 - Exhaust
List of Figures Figure 1: Figure 2: Figure 3: Figure 4: Figure 5: Figure 6: Figure 7: Figure 8:
Thrust Reverser Doors .........................................78-3 Thrust Reverser Operation....................................78-5 Thrust Reverser Component Location ..................78-7 Translating Sleeves and Sliders............................78-9 Blocker Door Fittings ...........................................78-11 Cascades ............................................................78-13 Thrust Reverser Actuation System......................78-15 Isolation Control Unit and Directional Control Unit Location.........................78-17 Figure 9: Thrust Reverser Actuation System Component Location ...........................................78-19 Figure 10: Isolation Control Unit and Directional Control Unit .......................................78-21 Figure 11: Locking Feedback Actuator................................78-23 Figure 12: Locking Actuator..................................................78-25 Figure 13: Locking Feedback Actuator and Locking Actuator Installation ...............................78-27 Figure 14: Track Lock Unit ...................................................78-29 Figure 15: Track Lock Valve.................................................78-31 Figure 16: Throttle Quadrant Assembly................................78-33 Figure 17: Locking Feedback Actuator Description..............78-35 Figure 18: Throttle Quadrant Assembly Thrust Lever Control............................................78-37 Figure 19: Thrust Reverser Control ......................................78-39 Figure 20: Thrust Reverser Indications ................................78-41 Figure 21: Thrust Reverser Locked ......................................78-43 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Figure 22: Isolation Control Unit Energized – Actuators Overstow ............................................ 78-45 Figure 23: Track Lock Units Released................................. 78-47 Figure 24: DCU Energized................................................... 78-49 Figure 25: Thrust Reversers Fully Deployed ....................... 78-51 Figure 26: Stow Command .................................................. 78-53 Figure 27: Thrust Reverser Stowed ..................................... 78-55 Figure 28: Thrust Reverser Actuating System Control Electrical Schematic............................... 78-57 Figure 29: Thrust Reverser Actuating System Logic ........... 78-59 Figure 30: Isolation Control Unit Deactivation...................... 78-63 Figure 31: Thrust Reverser Latch Beam Lockout Pins ........ 78-65 Figure 32: Manually Deploy the Thrust Reverser Sleeves .................................... 78-67 Figure 33: Thrust Reverser Cycling Test ............................. 78-69
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-ii
CS130_M5_78_ExhaustLOF.fm
78 - Exhaust
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CS130_M5_78_ExhaustLOF.fm
EXHAUST - CHAPTER BREAKDOWN EXHAUST CHAPTER BREAKDOWN
Thrust Reverser Structure
1 Thrust Reverser Actuation and Control
2
Thrust Reverser Indications
3
78 - Exhaust 78-36 Thrust Reverser Structure
78-36 THRUST REVERSER STRUCTURE GENERAL DESCRIPTION The PW1500G engine has a cascade-type thrust reverser (T/R). Two thrust reverser doors make up the thrust reverser structure. The thrust reverser doors enclose the engine core section and provide the fan bypass airflow path for forward thrust. The composite inner fixed structure (IFS) provides ventilation, fire suppression, and noise attenuation of the engine core. A torque box, mounted on the leading edge of the IFS, provides the mounting points for the actuators, the cascades, the blocker doors and drag links. The outer section of the thrust reverser door houses the translating sleeves. The translating sleeves deploy during landing to direct fan airflow forward and through the cascades to slow the aircraft. Each thrust reverser door is attached to the pylon at the hinge beam. The latch beam has latches that secure the doors closed.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Hinge Beam
Blocker Door
Cascades Torque Box
Inner Fixed Structure Latch Beam
NOTE Thrust reverser shown in deployed position.
CS1_CS3_7833_001
Translating Sleeve
Figure 1: Thrust Reverser Doors Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-3
CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Thrust Reverser Operation The inner fixed structure (IFS) and the translating sleeve form the thrust reverser (T/R) door. Each T/R door has five composite blocker doors mounted on the translating sleeve. The upper and lower doors are unique in size and cannot be installed at any other position. The blocker doors lay flush with the translating sleeve when in the stowed position. When the T/R is deployed, the drag links pull the blocker doors into the T/R door. This blocks the fan air from being directed out the exhaust. The airflow is directed in a controlled manner through the cascades in a forward and outward direction. The forward component of the redirected airflow provides the reverse thrust action which aids in slowing the aircraft. The outward component of the airflow is designed to ensure the reverse thrust air is not redirected into the engine inlet, or against the fuselage and flight control surfaces.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Upper Blocker Door
Drag Links
Drag Links
Translating Sleeve Inner Fixed Structure
BLOCKER DOORS IN STOWED POSITION
Lower Blocker Door BLOCKER DOORS IN DEPLOYED POSITION
CS1_CS3_7833_004 S3 7833 004
Blocker Doors
Figure 2: Thrust Reverser Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_78_L2_TRS_General.fm
78 - Exhaust 78-36 Thrust Reverser Structure
COMPONENT LOCATION
TRANSLATING SLEEVE
The thrust reverser structure has the following components:
The translating sleeves are mounted on the IFS at the hinge beam and the latch beam.
•
-
Inner fixed structure (IFS)
-
Hinge beam
Blocker Doors
-
Latch beam
The 10 blocker doors are mounted on the translating sleeves.
-
Torque box
-
Cascades
Translating sleeve -
Blocker doors
INNER FIXED STRUCTURE The inner fixed structure (IFS) encloses the engine core section. Hinge Beam The hinge beams are located on the top of the IFS. Latch Beam The latch beams are located at the bottom of the IFS. Torque Box The torque box is located at the front of the IFS. Cascades The cascades are mounted on the torque box.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_ComponentLocation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Translating Sleeve
Hinge Beam
Cascades
Torque Box Inner Fixed Structure
Latches
Latch Beam
CS1_CS3_7833_002
Blocker Doors
Figure 3: Thrust Reverser Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
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CS130_78_L2_TRS_ComponentLocation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
COMPONENT INFORMATION TRANSLATING SLEEVES The translating sleeves deploy aft guided by the primary and secondary sliders that move inside the primary and secondary Teflon coated slider tracks, which are attached to the inner fixed structure (IFS). Leading edge bumpers rest on lands on the torque box in order to maintain the aerodynamic contour of the translating sleeve when stowed. The forward section of the inner sleeve have cutouts in which the blocker doors rest flush to provide a streamlined fan bypass air flow.
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TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Primary Slider Track Secondary Slider Track Secondary Slider Blocker Doors Primary Slider
Lands
Leading Edge Bumper
Secondary Slider Track Torque Box
Secondary Slider
INTERIOR VIEW OF TRANSLATING SLEEVE
CS1_CS3_7833_003
Primary Slider
Primary Slider Track
Figure 4: Translating Sleeves and Sliders Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-9
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
BLOCKER DOORS The blocker doors have hinge fittings that mate with clevises attached to the inner wall of the translating sleeve. The doors rotate on these hinge fittings when the sleeves translate rearward. Drag links aid in rotating the blocker doors into the T/R door. The drag links are fixed-length aluminum rods with an airfoil cross-section. Teflon coated spherical bearings on both ends of the drag links ensure smooth movement. An aluminum drag link housing bracket built into the blocker door provides the attachment point for the drag link. A leaf spring mechanism absorbs drag link forces on the blocker door. Rest pads, mounted on supports, contact the blocker doors to absorb vibrations of the doors when stowed. Each blocker door has drain holes to prevent water accumulation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-10
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Blocker Door Seal Clevis
Hinge Fitting Leaf Spring
Drain Holes
Drag Link Housing
Drag Link
Clevis Hinge Fitting
CS1_CS3_7833_005
Spherical Bearings Rest Pad
Figure 5: Blocker Door Fittings Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-11
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
CASCADES The cascades are made up of sixteen individual carbon fiber cascade boxes, designed to turn fan airflow into forward and sideways directions. The cascade boxes are designed to provide specific airflow for their location on the engine. The cascade boxes have fastener patterns that ensure they can not be installed incorrectly. Each cascade box mounts between the torque box and the aft cascade ring. Flow blockers are installed at the hinge beam and latch beam to prevent the airflow from leaking between the cascade boxes and the beams.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-12
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-36 Thrust Reverser Structure
Flow Blocker
1
2
16 3
4
15
5
14
Aft Cascade Ring 6
13
7 12
Torque Box
8
10
9 CASCADE BOX LOCATION
CS1_CS3_7833_006
11
Figure 6: Cascades Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-13
CS130_78_L2_TRS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
78-31 THRUST REVERSER ACTUATION SYSTEM GENERAL DESCRIPTION The reverse actuation system has four hydraulic actuators, two track lock units, an isolation control unit (ICU), and a directional control unit (DCU). The thrust reverser actuation system (TRAS) is controlled through the electronic engine control (EEC), and the control and distribution cabinets (CDCs). The thrust reverser actuation system operates when the aircraft is on the ground. It is controlled by moving the thrust levers on the throttle quadrant assembly (TQA) to the reverse idle position. Each thrust reverser door has a translating sleeve that is operated by a locking feedback actuator and a locking actuator. The operating of the two actuators is synchronized to ensure proper movement of the translating sleeve. The locking feedback actuator has a linear variable differential transformer (LVDT) for position feedback. The actuators have an internal hydraulically released mechanical lock. A proximity switch monitors the lock status of each actuator. Two track lock units lock the translating sleeve when the thrust reverser is stowed. They protect against inadvertent thrust reverser deployment. Each track lock unit is controlled by a track lock solenoid valve. The track lock solenoid valve is controlled by the CDC. The track lock unit releases when the translating sleeve is ready to be deployed. The track lock unit relocks the translating sleeve once it stows. The thrust reverser actuation system (TRAS) has an isolation control unit (ICU) that isolates the system from supply pressure when the system is non-operational and a directional control unit (DCU) that controls the directional functions of the thrust reverser translating sleeves. The ICU and DCU have solenoid valves that are controlled by the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
To allow for dispatch of inoperative thrust reversers (T/Rs), and to provide the additional safety during maintenance activities, a manual inhibit lever is installed on the ICU to isolate the thrust reverser actuation system (TRAS). The DCU has a directional spool valve operated by a directional control solenoid valve that controls the unlock, deploy, stow and lock sequence when commanded. The thrust reverser (T/R) is deployed by moving the thrust levers from the forward thrust position to the idle position and then to the reverse idle position. Once the translating sleeves have reached 85% of full travel, the thrust levers can be moved to the maximum reverse thrust position. The EEC provides a higher maximum reverse thrust for a rejected takeoff than for a rollout after landing. The T/Rs are stowed by returning the thrust levers to the idle position. The thrust levers can not be moved beyond the idle position and engine thrust is limited until the thrust reversers are fully stowed. During the deploy sequence, a white REV icon is displayed on EICAS. When the thrust reverser (T/R) is fully deployed, the REV icon turns green. When the T/Rs are stowed, the white REV icon is displayed until the T/R is fully locked. The locking feedback actuator, locking actuator, and track lock unit provide mechanical protection against an inadvertent thrust reverser (T/R) deployment. The EEC controlled ICU and DCU, and the CDC controlled track lock solenoid valves provide electrical protection against an inadvertent thrust reverser (T/R) deployment. The EEC monitors the T/R position through the locking feedback actuator LVDTs, lock status, and the track lock unit lock status. If a thrust reverser inadvertently deploys, the EEC reduces the engine thrust to idle.
TECHNICAL TRAINING MANUAL
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78-14
CS130_78_L2_TRAS_General.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND
ISOLATION CONTROL UNIT
REV
Air/Gnd
TRACK LOCK VALVE
ELECTRONIC ENGINE CONTROL Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
CDC
DMC 1 and DMC 2
SOL
CDC DMC LVDT SOL
Pressure Return Control and Distribution Cabinet Data Concentrator Unit Module Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor ARINC 429
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_032
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 7: Thrust Reverser Actuation System Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRAS_General.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
COMPONENT LOCATION The thrust reverser actuating system has the following components: •
Isolation control unit (ICU)
•
Directional control unit (DCU)
•
Locking feedback actuators (Refer to Figure 9)
•
Locking actuators (Refer to Figure 9)
•
Flexible drive shaft/deploy tube (Refer to Figure 9)
•
Manual drive units (Refer to Figure 9)
•
Track lock valves (Refer to Figure 9)
•
Track lock units (Refer to Figure 9)
ISOLATION CONTROL UNIT The isolation control unit (ICU) is located in the aft pylon section and is accessible from access panels on each side of the pylon. DIRECTIONAL CONTROL UNIT The directional control unit (DCU) is located in the forward pylon section, above the engine core section area.
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CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
A
B
B DIRECTIONAL CONTROL UNIT
CS1_CS3_7832_005
A ISOLATION CONTROL UNIT
Figure 8: Isolation Control Unit and Directional Control Unit Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-17
CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING FEEDBACK ACTUATORS The locking feedback actuators are located on each thrust reverser door upper section. LOCKING ACTUATORS The locking actuators are located on each thrust reverser door lower section. FLEXIBLE DRIVE SHAFT/DEPLOY TUBES The flexible drive shaft/deploy tubes are located on the torque box between the locking actuator, and the locking feedback actuator. MANUAL DRIVE UNITS The manual drive units are located on each of the locking actuators. TRACK LOCK VALVES The track lock valves are located on each of the thrust reverser door lower section, near the latch beam. TRACK LOCK UNITS The track lock units are located on each of the thrust reverser latch beam.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-18
CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Locking Feedback Actuator
Return Deploy Stow Flexible Drive Shaft/Deploy Tubes
Locking Actuator
Track Lock Valves
Manual Drive Units
Locking Actuator
CS1_CS3_7832_002
Hydraulic Line
Track Lock Units
Figure 9: Thrust Reverser Actuation System Component Location Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-19
CS130_78_L2_TRAS_ComponentLocation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
COMPONENT INFORMATION ISOLATION CONTROL UNIT The isolation control unit (ICU) isolates the thrust reverser actuating system (TRAS) from the aircraft hydraulic system. The ICU has a dual-channel solenoid controlled by either electronic engine control (EEC) channel. The solenoid operates a pilot valve that supplies hydraulic pressure to an internal isolation spool valve. The isolation spool valve opens to allow hydraulic pressure from the aircraft hydraulic system to enter the TRAS. To allow for dispatch of an inoperative thrust reverser (T/R) and to provide additional safety during ground maintenance activities, a manual inhibit lever is installed to prevent the isolation spool valve from operating. An inhibit pin stored on the ICU locks the manual inhibit lever in the inhibit position. The position of the inhibit lever is monitored by dual-channel inhibit proximity sensors. DIRECTIONAL CONTROL UNIT The directional control unit (DCU) has a dual-channel solenoid controlled by either EEC channel. The solenoid operates a directional control valve that supplies hydraulic pressure to an internal directional spool valve. The directional spool valve provides hydraulic pressure to the actuators and the track lock valves to control unlocking, deploying, stowing, and locking of the thrust reversers (T/Rs). A dual-channel proximity pressure sensor provides the status of the hydraulic pressure at the outlet of the ICU to each channel of the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Dual-Channel Solenoid
Dual-Channel Inhibit Proximity Sensors Inhibit Pin
ISOLATION CONTROL UNIT
Manual Inhibit Lever
Dual-Channel Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT
CS1_CS3_7832_034
Dual-Channel Solenoid
Figure 10: Isolation Control Unit and Directional Control Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-21
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING FEEDBACK ACTUATOR The locking feedback actuator consists of a jackscrew that drives the translating sleeve. The jackscrew is hydraulically operated. The stow tube supplies hydraulic fluid to stow the translating sleeve. The rod end fitting at the aft end of the actuator provides a connection point to the translating sleeve. The actuator has an internal lock that maintains the actuator in the stowed position when hydraulic pressure is not applied. The lock releases when hydraulic pressure is applied. This allows the translating sleeve to deploy. When the translating sleeve stows, the lock re-engages. A proximity sensor monitors the lock and provides a stowed signal to the electronic engine control (EEC). The actuator has a manual lock that mechanically unlocks the actuator to allow the translating sleeve to be manually deployed. The locking feedback actuator connects to the locking actuator through a flexible drive shaft and deploy tube. The flexible drive shaft synchronizes the movement of the two actuators. The deploy tube supplies hydraulic fluid to the locking actuator. A linear variable differential transformer (LVDT), mounted on the forward end of the actuator, provides position feedback of the translating sleeves to the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-22
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Rod End
Stow Tube
Manual Lock
Jackscrew Shaft
Proximity Sensor Dual-Channel Linear Variable Differential Transformer
Flexible Drive Shaft
CS1_CS3_7832_033
Deploy Tube
Figure 11: Locking Feedback Actuator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-23
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING ACTUATOR The locking actuator is similar in construction to the locking feedback actuator except it does not have an LVDT and its position is not monitored. The actuator has an internal lock that maintains the actuator in the stowed position when hydraulic pressure is not applied. The lock releases when hydraulic pressure is applied. This allows the translating sleeve to deploy. When the translating sleeve stows, the lock re-engages. A proximity sensor monitors the lock and provides a stowed signal to the EEC. The actuator has a manual lock that mechanically unlocks the actuator to allow the translating sleeve to be manually deployed. The locking actuator connects to the locking feedback actuator through a flexible drive shaft and deploy tube. The flexible drive shaft synchronizes the movement of the two actuators. The deploy tube carries hydraulic fluid to the actuator. A manual drive unit (MDU) mates with the actuator flexible drive shaft. This rotational input drives the locking actuator through the flex drive shaft, and the locking feedback actuator. A torque limiter in the MDU prevents over torquing of the actuator during manual operation.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Flexible Drive Shaft
Deploy Tube
CS1_CS3_7832_009 32_009
Manual Lock
Proximity Sensor Manual Drive Unit
Figure 12: Locking Actuator Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-25
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LOCKING FEEDBACK ACTUATOR AND LOCKING ACTUATOR INSTALLATION The locking feedback actuator and locking actuators are mounted in gimbals on the torque box. The gimbal allow the actuator to pivot as the translating sleeve deploys. This prevents the translating sleeve from jamming. The rod end fitting at the aft end of the actuator is attached to the translating sleeve. A panel provides access to the rod end fitting.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-26
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Rod End Fitting
A
Translating Sleeve
B
A ROD END OF ACTUATOR SHOWN WITH ACCESS PANEL REMOVED
Gimbal
B
HYDRAULIC LOCKING FEEDBACK ACTUATOR
CS1_CS3_7832_006
Torque Box
Figure 13: Locking Feedback Actuator and Locking Actuator Installation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-27
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK UNIT The track lock unit is installed on the latch beam of each thrust reverser door. The track lock unit is operated by hydraulic pressure supplied through the track lock valves. The track lock unit has a spring-loaded bolt that protrudes into the path of the translating sleeve. In the event of inadvertent deployment, the track lock unit prevents the translating sleeve from deploying. The bolt retracts when hydraulic pressure is supplied. Two proximity sensors monitor the position of the track lock unit bolt and provide the lock status to the EEC. An unlock lever is used to retract the bolt when the translating sleeve is manually deployed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-28
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Track Lock Bolt
Track Lock Unit
Proximity Sensors
Translating Sleeve
CS1_CS3_7832_012
Unlocking Lever
Figure 14: Track Lock Unit Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-29
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK VALVE The track lock valve is a solenoid powered valve used to supply hydraulic pressure to unlock the track lock unit. When the translating sleeve deploys, the track lock valve solenoid stays energized. This allows the track lock unit to remain unlocked until the thrust reverser is returned to the stowed position.
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78-30
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CS1_CS3_7832_015
Solenoid Valve
Figure 15: Track Lock Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-31
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THROTTLE QUADRANT ASSEMBLY The throttle quadrant assembly (TQA) provides thrust lever position to the EEC. The EEC calculates the engine power required between REV and the MAX REV position 19.5° aft of IDLE. Each thrust lever has two thrust reverser stow/deploy switches that provide thrust lever positions to the control and distribution cabinet (CDC). These switches are activated by cams at 1.85° aft of thrust lever idle position. The TQA baulk mechanism prevents the thrust levers from being moved beyond 7.5° aft of IDLE until the thrust reverser translating sleeves are fully deployed. The baulk mechanism prevents the thrust levers from moving 2.5° forward of idle position until the translating sleeves are fully stowed.
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78-32
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Forward Thrust Baulk Release
IDLE
Reverse Switch Activation REV Idle Detent
1.85°
4.6°
2.5°
Reverse Thrust Baulk Release
Baulk Mechanism
MAX REV 7.5° 19.5°
Thrust Reverser Switches
Cams
CS1_CS3_7600_016
THROTTLE QUADRANT ASSEMBLY
Figure 16: Throttle Quadrant Assembly Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-33
CS130_78_L2_TRAS_ComponentInformation.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
DETAILED COMPONENT INFORMATION LOCKING FEEDBACK ACTUATOR AND LOCKING ACTUATOR When hydraulic pressure is supplied from the isolation control unit (ICU), it is routed through the stow tube, to the stow port of the actuator, keeping the actuator stowed and locked. When the thrust reverse is selected, hydraulic pressure, supplied through the deploy port is directed to the deploy side of the actuator. The hydraulic pressure pushes the lock sleeve clear of the lock tine, allowing the jackscrew to rotate. The hydraulic pressure acts on the internal side of the jackscrew. Both sides of the actuator are now pressurized. Since the jackscrew surface area is greater than the stow side of the actuator piston, the actuator deploys. When the thrust reversers (T/Rs) are commanded to stow, the deploy side of the actuator is directed to the return port and the hydraulic pressure on the stow side drives the actuator to the stow position. While deploying or stowing, the jackscrew worm wheel turns the flexible drive shaft worm gear. The flexible drive shaft worm gear is connected to the flexible drive shaft that interconnects the locking feedback actuator to the locking actuator. This allows the actuators to maintain synchronization during the stow or deploy cycle. The manual lock is used to move the lock sleeve when the translating sleeve is manually deployed. The LVDT on the locking feedback actuator is integral to the actuator and is not a line replaceable unit (LRU). The LVDT has an internal self-rigging mechanism that operates on the first deploy and stow cycle. The self-rigging mechanism establishes the stowed and locked position for the actuator. All position feedback signals to the engine electronic control (EEC) are referenced from the actuator stowed and locked position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-34
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Stow Tube
Manual Lock
Jack Screw Shaft
Dual-Channel Linear Variable Differential Transformer
Flexible Drive Shaftr
A
Gimbal Manual Lock Lock Sleeve
Actuator Piston
Worm Wheel
Deploy Port
Jackscrew
B Flexible Drive Shaft Input
Stow Hydraulic Pressure
A
Deploy Hydraulic Pressure
Lock Return Tine Port
Worm Wheel
Stow Port
CS1_CS3_7832_007
B
Figure 17: Locking Feedback Actuator Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-35
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THROTTLE QUADRANT ASSEMBLY THRUST LEVER CONTROL The gate plate, mounted on the TQA, has end stops at the IDLE, MAX and MAX REV positions. A gate pin on the thrust reverse link contacts the end stops to limit the movement of the thrust lever. An IDLE stop prevents the thrust lever from moving into the thrust reverse range. Thrust reverser finger lifts on each thrust lever raise the gate pin clear of the IDLE stop. When the gate pin is clear of the IDLE stop, the thrust lever can be moved into the thrust reverse range. The thrust lever can be returned to the forward thrust range without using the finger lifts. When the baulk solenoid energizes, it positions the baulk mechanism rocker arms to lock the thrust lever in the idle range until the thrust reverser is either fully deployed or fully stowed. If required, the interlock can be overridden with a force of approximately 25 lb.
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78-36
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Baulk Mechanism
Thrust Lever
B Thrust Reverser Finger Lifts
Rocker Arms
A
Thrust Reverser Link
IDLE Stop MAX Stop
Gate Pin
Gate Plate
A THRUST REVERSER BAULK MECHANISM (PLATE REMOVED FOR CLARITY)
B THRUST LEVER (COVER REMOVED FOR CLARITY)
CS1_CS3_7600_005 S1 CS3 7600 005
Baulk Solenoid
MAX REV Stop
Figure 18: Throttle Quadrant Assembly Thrust Lever Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-37
CS130_78_L3_TRAS_DetailedComponentInfo.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CONTROLS AND INDICATIONS CONTROLS The thrust reversers (T/Rs) are controlled from the throttle quadrant assembly (TQA). The thrust levers are moved into the reverse thrust range using the finger lifts located on the thrust levers. The thrust lever is used to control the amount of thrust required when the T/R has deployed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-38
CS130_78_L2_TRAS_ControlsIndications.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CS1_CS3_7831_001
Thrust Reverser Finger Lifts
Figure 19: Thrust Reverser Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-39
CS130_78_L2_TRAS_ControlsIndications.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
INDICATIONS The thrust reversers position is displayed on the engine indication and crew alerting system (EICAS) N1 gauge. The REV icon shows the status of the thrust reverser. •
A green REV icon appears in the N1 analog gauge when the thrust reversers (T/Rs) are fully deployed
•
A white icon indicates the thrust reversers are in transition
•
A amber REV icon indicates a T/R fault when the aircraft is on the ground, the thrust reverser is not manually inhibited, and any one of the following conditions occurs: -
Inadvertent thrust reverser deployment
-
Thrust reverser translating sleeve position feedback failed on either translating sleeve
-
Thrust reverser is jammed
-
Thrust reverser inability to deploy
-
Thrust reverser inadvertently pressurized
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CS130_78_L2_TRAS_ControlsIndications.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
REV
N1
70.4 REV TABLE 1 EICAS FLAG Symbol
688
692
REV
Aircraft is on ground, the thrust reverser is not manually inhibited, and any one of the following conditions: Ɣ7KUXVWUHYHUVHULQDGYHUWHQWGHSOR\PHQW Ɣ7KUXVWUHYHUVHUWUDQVODWLQJVOHHYHSRVLWLRQIHHGEDFN failed on either translating sleeve Ɣ7KUXVWUHYHUVHULVMDPPHG Ɣ7KUXVWUHYHUVHULQDELOLW\WRGHSOR\ Ɣ7KUXVWUHYHUVHULQDGYHUWHQWO\SUHVVXUL]HG
REV
On ground: thrust reverser in transition.
REV
On ground: thrust reverser is deployed.
EGT
87.5 1490 105 110
N2 FF (PPH) OIL TEMP OIL PRESS
87.5 1495 104 110
Description
NOTE The thrust reverser icon has priority over the vibration icon on the ground.
CS1_CS3_7837_005
70.5
Figure 20: Thrust Reverser Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-41
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78 - Exhaust 78-31 Thrust Reverser Actuation System
DETAILED DESCRIPTION THRUST REVERSER LOCKED When the thrust reverser (T/R) is stowed and the thrust levers are in forward thrust, the isolation control unit (ICU) and directional control unit (DCU) isolate hydraulic pressure from the thrust reverser system. The ICU isolation control valve control solenoid is de-energized. The directional control unit (DCU) pressure proximity sensors monitor the hydraulic pressure in the thrust reverser actuating system. The track lock valve solenoids are also de-energized and the track lock units are in the locked position. The track lock proximity sensors monitor the track lock units status. With the DCU directional control valve solenoid de-energized, the actuator is locked in the stowed position. The locking actuator proximity sensors monitor the lock mechanism position to ensure the thrust reverser is stowed and locked. The feedback actuator LVDTs provide position feedback as a secondary indication of the actuator position.
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78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND
RH Track Lock Position LH Track Lock Position
SOL
LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Linear Variable Differential Transformer Solenoid Proximity Sensor
TRACK LOCK VALVE
ICV Command ICU Pressure
DIRECTIONAL CONTROL UNIT
SOL
ISOLATION CONTROL UNIT Isolation Control Valve
Pressure Proximity Sensor
SOL
Directional Control Valve
Track Lock Unit
LOCKING ACTUATOR LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
NOTE Left engine reverser shown.
Track Lock Unit CS1_CS3_7832_018
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 21: Thrust Reverser Locked Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-43
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ISOLATION CONTROL UNIT ENERGIZED - ACTUATORS OVERSTOW When reverse thrust is selected, the electronic engine control (EEC) energizes the ICU isolation valve control solenoid. The hydraulic pressure is directed through the isolation spool valve to the DCU. The dual pressure proximity sensor, installed on the DCU, signals the EEC that hydraulic pressure is available from the ICU. Hydraulic pressure is supplied to the stow side of the locking feedback actuators and the locking actuators. The actuators over stow to release the locking mechanisms. The proximity sensors on the actuators indicate the locking mechanism is in the locked position. Hydraulic pressure is also supplied to the track lock valves. The track lock units remain locked. The track lock unit proximity sensors send the locked status to the EEC.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-44
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND SOL
LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Linear Variable Differential Transformer Solenoid Proximity Sensor
RH Track Lock Position
TRACK LOCK VALVE
LH Track Lock Position ICV Command ICU Pressure
ISOLATION CONTROL UNIT
DIRECTIONAL CONTROL UNIT
SOL
SOL
Isolation Control Valve
Pressure Proximity Sensor
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
NOTE Left engine reverser shown.
Track Lock Unit CS1_CS3_7832_019
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 22: Isolation Control Unit Energized – Actuators Overstow Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-45
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK UNITS RELEASED With the WOW logic satisfied, the control and distribution cabinet (CDC) energizes the track lock valve solenoids open both track lock valves. The hydraulic pressure is now supplied to the track lock unit and the track lock unit bolt retracts. The track lock unit proximity sensors signal the EEC indicating that the track lock unit bolts are clear of the translating sleeve and the thrust reverser is ready to deploy.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-46
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CDC TRACK LOCK VALVE COMMAND
LEGEND
RH Track Lock Position LH Track Lock Position
TRACK LOCK VALVE
SOL
CDC LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
ICV Command ICU Pressure
ISOLATION CONTROL UNIT
DIRECTIONAL CONTROL UNIT
SOL
SOL
Isolation Control Valve
Pressure Proximity Sensor
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_020
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 23: Track Lock Units Released Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-47
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
DCU ENERGIZED The EEC energizes the DCV directional control valve solenoid. This positions the directional spool valve to pressurize the actuators deploy side to completely unlock the actuators. The actuator proximity sensors and LVDTs provide the unlock status to the EEC and the translating sleeves start to deploy. The track lock valves solenoids remain energized during the deploy sequence. A white REV icon is displayed on the EICAS N1 gauge.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-48
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ELECTRONIC ENGINE CONTROL
LEGEND
ISOLATION CONTROL UNIT
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position LH Sleeve 95% Deployed RH Sleeve 95% Deployed
TRACK LOCK VALVE
SOL
CDC LVDT SOL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
CDC TRACK LOCK VALVE COMMAND
REV
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_021
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 24: DCU Energized Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-49
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSERS FULLY DEPLOYED The translating sleeve position is sent to the EEC by the locking feedback actuator LVDTs. When the translating sleeves reach 85% of full travel, the thrust levers may be moved to the full reverse thrust. When the translating sleeves reach 95% of full travel, the green REV icon is shown on the N1 gauge.
NOTE Both sides of the actuators are pressurized, however the greater actuator piston area on the deploy side ensures the actuators deploy.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-50
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ELECTRONIC ENGINE CONTROL
LEGEND
ISOLATION CONTROL UNIT
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position LH Sleeve 95% Deployed RH Sleeve 95% Deployed
TRACK LOCK VALVE
SOL
CDC LVDT SOL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
CDC TRACK LOCK VALVE COMMAND
REV
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_022
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 25: Thrust Reversers Fully Deployed Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-51
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
STOW COMMAND When the throttle moves from reverse idle to forward idle, the EEC de-energizes the DCU control solenoid. The directional spool valve repositions and the deploy side of the actuators are connected to the return side of the hydraulic system. Since hydraulic pressure always remain on the stow side of the actuators, the translating sleeves stow. The locking feedback actuator LVDTs feedback to the EEC changes the REV icon from green to white.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-52
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
ELECTRONIC ENGINE CONTROL
LEGEND
ISOLATION CONTROL UNIT
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT SOL
SOL
Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position LH Sleeve 95% Deployed RH Sleeve 95% Deployed
TRACK LOCK VALVE
SOL
CDC LVDT SOL
Pressure Return Control and Distribution Cabinet Linear Variable Differential Transformer Solenoid Proximity Sensor
CDC TRACK LOCK VALVE COMMAND
REV
Track Lock Unit
LOCKING ACTUATOR
Directional Control Valve
LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_023
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 26: Stow Command Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-53
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER STOWED The track lock valve solenoids de-energize, and the track lock units relock. When the actuator proximity sensors indicate the actuators are in the stowed position, the EEC de-energize the track lock valve solenoids. The spring-loaded track lock unit bolts extend, locking the translating sleeves. The track lock unit proximity sensors indicate the track lock unit lock status to the EEC. With the confirmation of the track lock units in the locked position, the EEC de-energizes the ICU control solenoid, isolating the trust reverser actuating system from the hydraulic system. The thrust reverser (T/R) hydraulic components are connected to the return line. The DCU pressure proximity sensors confirm the pressure is not available to the thrust reverser actuating system (TRAS). The REV indication is removed from the N1 gauge.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-54
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LEGEND
Pressure Proximity Sensor
DIRECTIONAL CONTROL UNIT
SOL
ISOLATION CONTROL UNIT Isolation Control Valve
ICV Command ICU Pressure DCV Command RH Track Lock Position LH Track Lock Position
SOL
Directional Control Valve
SOL
LVDT SOL
ELECTRONIC ENGINE CONTROL
Pressure Return Linear Variable Differential Transformer Solenoid Proximity Sensor
TRACK LOCK VALVE
Track Lock Unit
LOCKING ACTUATOR LOCKING FEEDBACK ACTUATOR
RIGHT
LVDT Inlet Screen
Hydraulic System No. 1
Isolation Spool Valve
Thrust Reverser Translating Sleeve
Directional Spool Valve
LOCKING FEEDBACK ACTUATOR LVDT
Return LEFT LOCKING ACTUATOR
Track Lock Unit
NOTE Left engine reverser shown.
CS1_CS3_7832_024
TRACK LOCK VALVE
SOL
MANUAL INHIBIT LEVER
Figure 27: Thrust Reverser Stowed Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-55
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER ACTUATING SYSTEM CONTROL ELECTRICAL SCHEMATIC
The track lock units have a proximity sensor to send a signal to the EEC when the track lock units are unlocked.
The thrust levers in the throttle quadrant assembly (TQA) provide the command signal to the EEC to control the thrust reverser actuating system (TRAS). When the thrust lever is moved to the REV IDLE position, the stow/deploy switches are made. The stow/deploy switches send a discrete to the thrust reverser lock solid-state power controller (SSPC) in control and distribution cabinet 1 (CDC 1).
The EEC controls the TQA baulk solenoid based on LVDTs and the stow and lock proximity sensors feedback received from the actuators. The EEC also monitors the DCU status. The LVDTs are dual-channel with one output wired to each channel of the EEC.
The baulk solenoid that prevents the thrust lever from moving towards the MAX REV range is powered by the TQA L Reverser Lock SSPC in CDC 3. Once the thrust reversers are deployed beyond 85%, the EEC provides a ground to energize the baulk solenoid. The EEC controls the ICU and DCU control solenoids. The ICU solenoid receives power from DC BUS 1 from CDC 1 on channel A and DC BUS 2 from CDC 2 on channel B. The combinational logic in the CDCs enables them when the stow/deploy switches are made, and one of the following conditions is true: •
Weight-on-wheels (WOW) signal from the landing gear and steering control unit (LGSCU)
•
Wheel spin signal from the brake data concentrator unit (BDCU)
The WOW signal from the LGSCU is hardwired into the CDCs, and sent through the data concentrator unit module cabinets (DMCs) to the EEC. The EEC provides the ground to energize the ICV solenoid. The DCU solenoid receives both power and a ground from the EEC. A dual channel pressure sensor provides the EEC with a signal when hydraulic pressure is flowing through the ICU and into the DCU. Once the ICU and DCU control solenoids become energized, the control valves are actuated which allows hydraulic pressure to flow to the actuators and track lock valves. The track lock valves are controlled by solenoids. These solenoids are energized by DC BUS 1 from CDC 1 when the stow/deploy switches are made and the aircraft is on the ground.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
Once the T/Rs are deployed beyond 85%, the EEC energizes the baulk solenoid. This allows the thrust lever to move into MAX REV range. When the thrust reversers deploy beyond 95%, the EEC sends a signal to the EICAS for indicating the green REV icon. Maximum reverse thrust is determined by the EEC. The EEC makes this determination using thrust lever angle (TLA). If the TLA was above 23° and then retarded to idle, the EEC determines that a rejected takeoff has occurred and sets the maximum reverse thrust level at 87.4% N1. During rollout after landing, the maximum reverse thrust is 80% N1. If the airspeed is below 55 kt and reverse thrust is still applied, the maximum reverse thrust N1 is reduced to reverse idle N1. If either thrust lever is moved in the reverse thrust range while the aircraft is in flight, a THROTTLE IN REVERSE caution message is displayed. Inadvertent operation of the thrust reverser (T/R) is detected by the EEC through monitoring the T/R position feedback and the lock status. If the T/R inadvertently deploys, the EEC commands the engine power to idle. An L(R) REVERSER FAIL caution message is displayed if the EEC detects faults that prevent system operation. On the ground, an amber REV icon is also displayed in the N1 indicator. If both locks on a translating sleeve indicate unlocked, a track lock indicates unlocked, or the TRAS is inadvertently pressurized, then a L(R) REVERSER UNLOCK caution message is displayed. The T/R may still be usable for landing. A dual-channel proximity sensor monitors the manual inhibit lever on the ICU. If the thrust reverser (T/R) is locked out, all other T/R CAS messages are inhibited and an L(R) REVERSER INHIBIT status message is displayed.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-56
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
TRACK LOCK UNIT
CDC 1-10-9 L Thrust Reverser Lock DC SSPC 5A BUS 1
DMC 1 and DMC 2
LGSCU
Logic: Advanced
TQA
CDC 3-4-15 TQA L Reverser Lock DC SSPC 3A BUS 1
LVDT
ACTUATOR Baulk Solenoid ACTUATOR
CDC 1-11-14 L Thrust Reverser ICV A DC SSPC 3A BUS 1
ACTUATOR
EEC
LVDT TRACK LOCK UNIT
Channel A
Logic: Combinational
TRACK LOCK SOLENOID VALVE
MANUAL INHIBIT
Proximity Switch
Channel A Channel B
Channel B
ICU
DCU
ICV ENABLE SOL
DCV ENABLE SOL
Channel A
Channel B
Channel B PRESSURE SENSOR
Return Isolation Control Valve
Hydraulic System 1
LEGEND
Channel A
Channel A
Channel B NOTE Left engine shown. Right engine similar.
Directional Control Valve
BDCU CDC DCU DCV DMC EEC ICU ICV LGSCU LVDT TQA
Pressure Return Brake Data Concentrator Unit Control and Distribution Cabinet Directional Control Unit Directional Control Valve Data Concentrator Unit Module Cabinet Electronic Engine Control Isolation Control Unit Isolation Control Valve Landing Gear and Steering Control Unit Liner Variable Differential Transformer Throttle Quadrant Assembly ARINC 429
CS1_CS3_7832_025
Logic: Combinational
ACTUATOR
Stow Deploy Switches
Logic: Always On
CDC 2-9-11 L Thrust Reverser ICV B DC SSPC 3A BUS 2
Proximity Switch
TRACK LOCK SOLENOID VALVE
BDCU
Figure 28: Thrust Reverser Actuating System Control Electrical Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-57
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER ACTUATING SYSTEM CONTROL LOGIC Isolation Control Valve Logic
Directional Control Valve Logic
The ICV logic requires an aircraft on the ground signal supplied by landing gear and steering control unit (LGSCU) 1 or LGSCU2 WOW signal, or an aircraft touchdown signal supplied by any 2 of the 4 wheel speed units (WSU) indicating a wheel speed of 45 kt.
The DCU logic requires an aircraft on the ground signal supplied by landing gear and steering control unit (LGSCU) 1 or LGSCU2 WOW signal or an aircraft touchdown signal supplied by any left WSU and any right WSU indicating a wheel speed of 45 kt.
The ICV energizes when:
The DCV energizes when:
•
Thrust reverse is selected
•
Thrust reverse is selected
•
No manual inhibit
•
No manual inhibit
•
No thrust reverser fault detected
•
No thrust reverser fault detected
•
Engine running
•
Track lock units unlocked
•
Engine running
The ICV also energizes on the ground when the engine is not running and the EEC thrust reverser cycling test is carried out through the onboard maintenance system (OMS). Track Lock Solenoid Control Logic Once the aircraft is on the ground, the landing gear and steering control unit (LGSCU) 1 or LGSCU 2 supply a WOW signal to the track lock logic. When the thrust levers are selected to the reverse thrust range beyond the -1.85° and both left stow/deploy switches 1 and 2 are actuated by the thrust lever, the track lock solenoid is energized When the thrust reverser is stowed, the track lock solenoids are energized for 11 seconds to provide enough time for the translating sleeve to stow.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The DCV also energizes on the ground when the engine is not running and the EEC thrust reverser cycling test is carried out through the onboard maintenance system (OMS). Baulk Solenoid Logic When thrust reverse is selected, each thrust lever is locked in the reverse idle position by the baulk mechanism. The thrust lever is locked until the thrust reverser translating sleeves have deployed far enough to safely increase thrust. When the directional control valve (DCV) has energized and the translating sleeves have deployed beyond 85%, the baulk solenoid energizes and releases the baulk mechanism. The thrust lever can be moved beyond the reverse idle position.
TECHNICAL TRAINING MANUAL
For Training Purposes Only
78-58
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
LGSCU WOW 1 LGSCU WOW 2 2 Out of 4 Wheel Speed Units > 45 Knots
Manual Inhibit No Fault Detected Thrust Reverse Selected
Isolation Control Unit Solenoid Energized
Engine Running Engine Off and EEC Thrust Reverser Cycling Test ISOLATION CONTROL UNIT SOLENOID LOGIC LGSCU WOW 1 LGSCU WOW 2
Track Lock Solenoid Energized Left Stow/Deploy Switch 1 Left Stow/Deploy Switch 2 TRACK LOCK SOLENOID LOGIC
1 of 2 Left Wheel Speed Units > 45 Knots 1 of 2 Right Wheel Speed Units > 45 Knots
LGSCU WOW 1 LGSCU WOW 2
Manual Inhibit No Fault Detected Thrust Reverse Selected Track Locks Unlocked
Directional Control Unit Solenoid Energized
Engine Running Engine Off and EEC Thrust Reverser Cycling Test DIRECTIONAL CONTROL UNIT SOLENOID LOGIC
Baulk Solenoid Energized
Left Translating Sleeve 85% Deployed LEGEND EEC LGSCU WOW
Electronic Engine Control Landing Gear and Steering Control Unit Weight-On-Wheels
Right Translating Sleeve 85% Deployed NOTE
BAULK SOLENOID LOGIC
Left engine shown. Right engine similar.
CS1_CS3_7832_026
Directional Control Unit Solenoid Energized
Figure 29: Thrust Reverser Actuating System Logic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-59
CS130_78_L3_TRAS_DetailedDescription.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
MONITORING AND TESTS The following page provides the crew alerting system (CAS) and INFO messages for the thrust reverser system.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-60
CS130_78_L2_TRAS_MonitoringTests.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
CAS MESSAGES Table 1: CAUTION Messages MESSAGE
LOGIC
L REVERSER FAIL
Left thrust reverser failed or not available. Icon is only shown if T/R has been commanded.
R REVERSER FAIL
Right thrust reverser failed or not available. Icon is only shown if T/R has been commanded.
L REVERSER UNLOCK
Multiple left reverser locks unlocked.
R REVERSER UNLOCK
Multiple right reverser locks unlocked.
THROTTLE IN REVERSE
Left or right thrust reverser selected in flight.
Table 2: ADVISORY Messages MESSAGE
LOGIC
L ENGINE FAULT Loss of redundant non-critical function for the left engine. R ENGINE FAULT Loss of redundant non-critical function for the right engine.
Table 3: STATUS Messages MESSAGE
LOGIC
L REVERSER INHIBIT
Left T/R manually inhibited as per maintenance procedure (crew awareness).
R REVERSER INHIBIT
Right T/R manually inhibited as per maintenance procedure (crew awareness).
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-61
CS130_78_L2_TRAS_MonitoringTests.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
PRACTICAL ASPECTS LOCKOUT THE THRUST REVERSERS FOR DISPATCH The T/Rs can be locked out for dispatch per the Aircraft Minimum Equipment List (MEL). The following is a synopsis of the Aircraft Maintenance Publication (AMP) procedure. Always consult the AMP before beginning any maintenance on the aircraft. There are two steps to deactivate the thrust reversers: •
Deactivate the isolation control unit (ICU)
•
Install a T/R lockout pin on the latch beam assembly
ICU DEACTIVATION The ICU is deactivated by using an inhibit lever located on the ICU, and installing a safety pin to maintain the lever in the inhibited position. The safety pin is stowed on the ICU when not required. Access to the ICU is through access panels on each side of the pylon.
WARNING THE THRUST REVERSER ACTUATION SYSTEM SHOULD BE DEACTIVATED USING THE ICU INHIBIT LEVER, BEFORE PERFORMING ANY MAINTENANCE. THE ACCIDENTAL OPERATION OF THE THRUST REVERSER CAN CAUSE INJURIES TO PERSONNEL, AND DAMAGE TO EQUIPMENT.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-62
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER OPERATIONAL
THRUST REVERSER LOCKED OUT
CS1_CS3_7836_001
REMOVE BEFORE FLIGHT
Lockout Pin Stowage
Figure 30: Isolation Control Unit Deactivation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-63
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Thrust Reverser Latch Beam Lockout Pin After locking out the ICU, lockout pins are installed on the latch beam assemblies to prevent the T/Rs from deploying. The lockout pins are stored on the thrust reverser door latch access panel. The pins are removed from their storage location and installed on the latch beam. The blanking plates that cover the lockout pin holes are stored on the latch access panel when the lockout pins are installed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-64
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
B
A Lockout Pin Blanking Plate
A
Lockout Pins
Latch Access Panel
NOTE Left side shown, right side similar.
B
CS1_CS3_7832_029
Translating Sleeve
Figure 31: Thrust Reverser Latch Beam Lockout Pins Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-65
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Manually Stow the Translating Sleeve
MANUALLY OPERATE THE TRANSLATING SLEEVES FOR MAINTENANCE The locking feedback actuators, locking actuators, and the track lock unit must be unlocked before manually operating the translating sleeves. Manually Deploy the Translating Sleeve Track Lock Unit Unlocking To unlock the track lock unit: •
Move the manual lock from the lock position to the unlock position
•
Install the unlock pin in the manual lockout hole.
Locking Feedback Actuator and Locking Actuator Unlocking To unlock the locking feedback actuator and locking actuator: •
Move the manual lock on the locking feedback actuator to the unlock position and install the unlock pin
•
Move the manual handle on the locking actuator to the unlock position and install the unlock pin
Translating Sleeve Deployment
Stow the Translating Sleeve Manually stow the translating sleeve as follows: •
Put the square drive tool in the manual drive unit
•
Turn the manual drive unit to stow the translating sleeve
•
Put the cover on the manual drive unit
Track Lock Unit Locking To lock the track lock unit: •
Remove the unlock pin from the manual lockout hole
•
Move the manual lock from the unlock position to the lock position
Locking Feedback Actuator and Locking Actuator Locking To lock the locking feedback actuator and locking actuator: •
Remove the unlock pin on the locking feedback actuator and move the manual lock back to the lock position
•
Remove the unlock pin on the locking actuator and move the manual lock back to the lock position
Manually deploy the translating sleeve as follows: •
Remove the cover and put the square drive tool in the manual drive unit
•
Turn the manual drive unit to extend the translating sleeve to the position necessary to do the maintenance
NOTE A maximum speed of 120 rpm is permitted for the manual drive unit. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-66
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Manual Drive Unit
Square Drive Tool Engagement
Locking Actuator
Lock Pin Manual Lock
Manual Lock
Proximity Sensor
Lock Pin
ACTUATOR MANUAL UNLOCKING FEATURE Manual Lockout Hole NOTE TRACK LOCK UNIT IN LOCKED POSITION
Locking actuator shown, locking feedback actuator similar.
CS1_CS3_7832_030
Cover
Figure 32: Manually Deploy the Thrust Reverser Sleeves Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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78-67
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
THRUST REVERSER CYCLING TEST The thrust reverser cycling test is used to detect any operational faults of the thrust reverser system including the electrical interfaces, mechanical linkage, and translating sleeve. This is done without an engine start.
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CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
MAINT MENU EEC1-A-THRUST REVERSER CYCLING TEST
1/8
*** WARNING *** THRUST REVERSER WILL DEPLOY AND STOW DURING THIS TEST ENSURE THERE ARE NO PERSONNEL NEAR THRUST REVERSER
PRECONDITIONS 1- REFER TO AMP TO PERFORM TASK XX-XX-XX 2- FADEC IS POWERED 3- THRUST REVERSER HYDRAULIC PRESSURE IS AVAILABLE 4- THRUST REVERSER IS NOT INHIBITED 5- THROTTLE LEVER IS SET TO FWD IDLE 6- PRESS "CONTINUE" TO GO TO NEXT PAGE
* NOTE *
YYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYYY
Continue
Write Maintenance Reports
Cancel
CS1_CS3_7321_026
TEST WILL TIMEOUT AFTER 5 MINUTES INACTIVITY
Figure 33: Thrust Reverser Cycling Test Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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78-69
CS130_78_L2_TRAS_PracticalAspects.fm
78 - Exhaust 78-31 Thrust Reverser Actuation System
Page Intentionally Left Blank
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TECHNICAL TRAINING MANUAL
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78-70
CS130_78_ILB.fm
ATA 79 - Oil
BD-500-1A10 BD-500-1A11
79 - Oil
Table of Contents 79-11 Storage .........................................................................79-2 General Description ..........................................................79-2 Practical Aspects .............................................................79-4 Oil Servicing ................................................................79-4 79-20 Distribution....................................................................79-6 General Description .........................................................79-6 Component Location ........................................................79-8 Lubrication and Scavenge Oil Pump............................79-8 Oil Control Module .......................................................79-8 Deoiler........................................................................79-10 Fuel/Oil Heat Exchanger Bypass Valve .....................79-10 Fuel/Oil Heat Exchanger............................................79-10 Variable Frequency Generator Oil/Oil Heat Exchanger .............................79-10 Air/Oil Heat Exchanger ..............................................79-10 Component Information ..................................................79-12 Lubrication and Scavenge Oil Pump .........................79-12 Oil Control Module .....................................................79-14 Deoiler........................................................................79-16 Main Oil Filter ............................................................79-18 Fan Drive Gear System Oil Pump..............................79-20 Detailed Component Information ....................................79-22 Journal Oil Shuttle Valve............................................79-22 Variable Oil Reduction Valve .....................................79-22 Detailed Description ........................................................79-24 Oil System Control .....................................................79-26
Detailed Component Information ................................... 79-34 Oil Debris Monitor ..................................................... 79-34 Controls and Indications ................................................ 79-36 EICAS ....................................................................... 79-36 Status ........................................................................ 79-36 Detailed Description ...................................................... 79-38 Oil Pressure Warning Limits...................................... 79-40 Monitoring and Tests ..................................................... 79-42 CAS Messages ......................................................... 79-43
79-30 Indicating ....................................................................79-30 General Description .......................................................79-30 Component Location ......................................................79-32 Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-i
CS130_M5_79_OilTOC.fm
79 - Oil
List of Figures Figure 1: Oil Storage ............................................................79-3 Figure 2: Oil Servicing ..........................................................79-5 Figure 3: Oil Distribution .......................................................79-7 Figure 4: Component Location Right Side............................79-9 Figure 5: Component Location Left Side ............................79-11 Figure 6: Lubrication and Scavenge Oil Pump ...................79-13 Figure 7: Oil Control Module...............................................79-15 Figure 8: Deoiler .................................................................79-17 Figure 9: Main Oil Filter ......................................................79-19 Figure 10: Fan Drive Gear System Oil Pump .......................79-21 Figure 11: Journal Oil Shuttle Valve and Variable Oil Reduction Valve...............................79-23 Figure 12: Distribution Schematic.........................................79-25 Figure 13: Oil System Control ..............................................79-27 Figure 14: Fan Drive Gear System Operation ......................79-29 Figure 15: Oil Indicating........................................................79-31 Figure 16: Oil System Indicating Components .....................79-33 Figure 17: Oil Debris Monitor................................................79-35 Figure 18: Oil Indications......................................................79-37 Figure 19: Oil Indicating Detailed Description ......................79-39 Figure 20: Oil Pressure Warning Limits................................79-41
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79-ii
CS130_M5_79_OilLOF.fm
79 - Oil
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TECHNICAL TRAINING MANUAL
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79-iii
CS130_M5_79_OilLOF.fm
OIL - CHAPTER BREAKDOWN OIL CHAPTER BREAKDOWN
Storage
1
Distribution
2
Monitoring and Tests
3
79 - Oil 79-11 Storage
79-11 STORAGE GENERAL DESCRIPTION A 27.3 L (28.8 qt) aluminum oil tank stores the oil and supplies it to the engine. The tank is wrapped in a stainless steel heatshield, held together by springs. The oil tank has an internal swirl-type deaerator that removes air from the returning oil. A pressure-regulating valve (PRV) maintains 12 psi head pressure in the tank. Excess pressure is vented to the deoiler in the main gearbox. The oil level can be monitored using a sight glass. A fill port has a hinged cap for oil servicing. The fill port strainer prevents foreign objects from entering the tank during servicing. In the event that the filler cap is incorrectly installed, a flapper valve prevents rapid loss of oil. A scupper drain collects oil spilt during servicing, and sends it to the drain mast. Oil can be drained through a drain plug in the bottom of the tank.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-2
CS130_79_L2_Storage_General.fm
79 - Oil 79-11 Storage
Oil Fill Port
Sight Glass
A
Pressurization Valve Scavenge Return Spring-Loaded Lock To Deoiler
Hinged Cap
To Lubrication and Scavenge Oil Pump
Flapper Valve A OIL TANK
B
B OIL FILL PORT
Scupper Drain
CS1_CS3_7911_002
Drain Plug
Figure 1: Oil Storage Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-3
CS130_79_L2_Storage_General.fm
79 - Oil 79-11 Storage
PRACTICAL ASPECTS
Continue to hold the flapper valve open. Add the recommended oil into the filler neck until no more oil can be added without overflow into the scupper drain.
OIL SERVICING The oil level can be checked and serviced by opening an access door on the right side of the core cowl fixed structure. The tank is equipped with a flapper valve to ensure the oil is not lost, should the cap remain open during flight. The sight glass is marked in both liters and quarts, and gauge markings indicate how much oil may be added.
WARNING A MINIMUM WAIT OF 5 MINUTES IS REQUIRED TO MAKE SURE THE OIL SYSTEM IS NOT PRESSURIZED PRIOR TO SERVICING THE OIL. FAILURE IN FOLLOWING THIS WARNING MAY RESULT IN PERSONAL INJURY.
Oil quantities are as follows: •
Fully serviced 21.8 L (23 qts)
•
Minimum 9.3 L (9.8 qts)
CAUTION
If the oil level is checked between 5 minutes and 60 minutes after engine shutdown, add oil to bring the oil level to the FULL mark. If the oil level is checked between 1 hour and 10 hours after engine shutdown, add oil only if the level is below the 3 L (3qt) mark on the sight glass. Do not add more than 3 L (3.1 qt). If the oil level is still not at the 3 L (3qt) mark, the engine must be run at idle until the oil is at the correct operating temperature. Shut down the engine and check the oil level. If required, add oil to bring the oil level to the FULL mark. If the oil level is checked more than 10 hours after engine shutdown, dry motor the engine until the engine oil pressure is stable. Check the oil quantity on the STATUS synoptic page. If the oil level is less than 15.1 L (16.0 US qts), add oil until the oil level is at the 3 L (3 qts) mark on the sight glass. Run the engine at ground idle until it is at operating temperature, then shutdown the engine. If required, add oil to bring the oil level to the FULL mark. To service the oil, lift the hinged cap and look into the oil tank to ensure the flapper is not stuck in a closed position. If the flapper is stuck, insert a small screwdriver through one of the holes in the oil tank inlet screen and open the oil tank flapper valve. Copyright © Bombardier Inc. CS130-21.07-06-00-121418
1. Check the oil level between 5 minutes and 1 hour after engine shutdown. Outside of this range, the oil level sight glass indication is not accurate. 2. Do not motor the engine before checking the oil level during the 1 hour to 10 hours after engine shutdown interval. Too much oil can be added and cause damage to the engine. 3. Do not damage the flapper valve when using a screwdriver to open it. 4. Use only engine oil specified in the service bulletin. The mixing of different brands of approved oils is not recommended, but is permitted within the limits specified in the service bulletin. The use of unapproved types or brands of oils is not permitted, and may cause damage to the engine. 5. Make sure the oil tank is completely closed. Improper closure and locking of the handle may result in oil exiting the tank, resulting in an inflight shutdown.
TECHNICAL TRAINING MANUAL
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79-4
CS130_79_L2_Storage_PracticalAspects.fm
79 - Oil 79-11 Storage
Spring-Loaded Lock Hinged Cap
A B
Flapper Valve
B OIL LEVEL SIGHT GLASS A
OIL FILL PORT
CS1_CS3_7911_003
OIL TANK
Figure 2: Oil Servicing Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-5
CS130_79_L2_Storage_PracticalAspects.fm
79 - Oil 79-20 Distribution
79-20 DISTRIBUTION GENERAL DESCRIPTION The main function of the engine oil system is to cool and lubricate the bearings, gearboxes, shafts and splines throughout the engine, and provide system status and fault indications for display on the engine indication and crew alerting system (EICAS). It also provides a thermal management function, where heat exchangers are used to cool the oil and heat the fuel. The system consists of the oil tank, pressure pump, oil control module (OCM), filters, heat exchangers, and associated plumbing. The plumbing connections on the engine core have been minimized by the use of the OCM, which is the central distribution point in the oil system. The dual-element main oil filter, installed on the OCM, has a primary filter and a primary filter bypass valve, which bypasses flow around the primary filter element in case of filter clogging. A secondary filter still continues to provide filtering in the event of a bypass. In addition to supplying oil pressure to the engine gearbox and bearings, the lubrication and scavenge oil pump (LSOP) also removes oil from the bearing cavities and gearboxes, and returns it to the tank. Pressure buildup in the compressor intermediate case causes the no. 3 bearing compartment and angle gearbox to experience increased pressure that must be released. The released air pressure, mixed with oil vapor, is called breather air. An external tube sends no. 3 bearing breather air to a deoiler unit in the main gearbox. The deoiler removes the oil droplets from the air. This air is sent overboard through the deoiler vent duct. Additional breather air comes from the oil tank. Oil tank pressure is released through the tank pressure valve when the tank pressure reaches the maximum limit The fan drive gear system (FDGS) has a dual-stage fan-driven oil pump that works with the journal oil shuttle valve (JOSV) to form an integrated Copyright © Bombardier Inc. CS130-21.07-06-00-121418
oil system, supplying oil to the FDGS journal bearings during normal, windmill, and negative-G conditions. A variable oil reduction valve (VORV) supplies pressurized oil from the engine oil system during normal operation. The VORV controls the amount of oil supplied based on engine power settings. The front bearing compartment contains two sumps that return oil back to the main oil tank from the components in the front bearing compartment. The sump, at the bottom of the front bearing compartment, also acts to supply oil to the journal bearings during windmill conditions. In normal operation the sump oil is scavenged and returned to the oil tank by the fan-driven oil pump. A secondary sump collects oil from the FDGS via an oil collection gutter. This secondary sump in the front compartment supplies oil during negative-G events. In normal operation, the fan-driven oil pump scavenges the oil and returns it to the oil tank. The thermal management system, consisting of three heat exchangers, ensures that engine oil and fuel temperatures are maintained within limits. The air/oil heat exchanger (AOHX) uses the fan air to cool the engine oil. If the AOHX becomes clogged, the pressure-relief valve (PRV) diverts oil directly to the variable frequency generator (VFG) oil/oil cooler. The fuel/oil heat exchanger (FOHX) is used to decrease oil temperature, and increase fuel temperature by transferring heat from the engine oil to the engine fuel. The FOHX has integral pressure-relief valves for both the fuel and oil sides of the heat exchanger. The fuel/oil heat exchanger bypass valve (FOHXBV) is a modulating valve used to control oil flow between the FOHX and AOHX based on fuel temperature. The variable frequency generator oil/oil heat exchanger (VFGOOHX) uses cooled engine oil to remove heat from the oil used by the VFG. An oil debris monitor and multiple chip collectors located on the LSOP provide system health monitoring.
TECHNICAL TRAINING MANUAL
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79-6
CS130_79_L2_Distribution_General.fm
79 - Oil 79-20 Distribution
LEGEND
OIL TANK
AGB Deoiler Overboard Breather No. 3 Bearing Breather
MAIN ENGINE BEARINGS 1,1.5,2,3,4,5,6 OCM
LUBRICATION AND SCAVENGE OIL PUMP
ODM
FDGS
Main Oil Filter
Fan Air
FDG VORV JOSV
AIR/OIL HEAT EXCHANGER
Fan Air Overboard
MGB
FUEL/OIL HEAT EXCHANGER BYPASS VALVE To Main Pump
To VFG
FUEL/OIL HEAT EXCHANGER OIL/OIL HEAT EXCHANGER
Journal Bearings
FDOP
From Boost Pump
Auxiliary Pump Stage
SECONDARY SUMP
Windmil Pump Stage
SUMP
From VFG
CS1_CS3_7920_003
AGB FDG FDGS FDOP JOSV MGB OCM ODM VFG VORV
FDGS Circuit Breather Air Fuel Pressure Scavenge VFG Oil System Supply Angle Gear Box Fan Drive Gear Fan Drive Gear System Fan-Driven Oil Pump Journal Oil Shutoff Valve Main Gear Box Oil Control Module Oil Debris Monitor Variable Frequency Generator Variable Oil Reduction Valve
Figure 3: Oil Distribution Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-7
CS130_79_L2_Distribution_General.fm
79 - Oil 79-20 Distribution
COMPONENT LOCATION The oil distribution system includes the following components: •
Lubrication and scavenge oil pump (LSOP)
•
Oil control module (OCM)
•
Deoiler
•
Main oil filter
•
Active oil damper valve (AODV)
•
Variable oil reduction valve (VORV)/journal oil shuttle valve (JOSV)
•
Fuel/oil heat exchanger bypass valve (FOHXBV) (Refer to figure 5)
•
Fuel/oil heat exchanger (FOHX) (Refer to figure 5)
•
Variable frequency generator oil/oil heat exchanger (VFGOOHX) (Refer to figure 5)
•
Air/oil heat exchanger (AOHX) (Refer to figure 5)
LUBRICATION AND SCAVENGE OIL PUMP Lubrication and scavenge oil pump (LSOP) is installed on the right front side of the main gearbox at the 5 o’clock position. OIL CONTROL MODULE Oil control module is mounted to the right side of the main gearbox.
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79-8
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
Active Oil Damper Valve
Oil Filter
Variable Oil Reduction Valve Lubrication and Scavenge Oil Pump
Journal Oil Shuttle Valve OIL CONTROL MODULE
CS1_CS3_7920_009 920_009
Deoiler
Figure 4: Component Location Right Side Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-9
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
DEOILER The deoiler is installed on the right side of the main gearbox. FUEL/OIL HEAT EXCHANGER BYPASS VALVE The fuel/oil heat exchanger bypass valve installed on the front side of the core engine at the 12 o’clock position. FUEL/OIL HEAT EXCHANGER The fuel/oil heat exchanger is installed on the left side of the core at the 11 o’clock position. VARIABLE FREQUENCY GENERATOR OIL/OIL HEAT EXCHANGER The variable frequency generator oil/oil heat exchanger is installed on the left side of the core engine at approximately the 10 o’clock position. AIR/OIL HEAT EXCHANGER The air/oil heat exchanger is installed on the left side of the core engine at the 10 o’clock position.
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79-10
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
A C
D
A
B
FUEL/OIL HEAT EXCHANGER
C
VFG OIL/OIL HEAT EXCHANGER
FUEL/OIL HEAT EXCHANGER BYPASS VALVE
D
AIR/OIL HEAT EXCHANGER
CS1_CS3_7920_008
B
Figure 5: Component Location Left Side Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-11
CS130_79_L2_Distribution_ComponentLocation.fm
79 - Oil 79-20 Distribution
COMPONENT INFORMATION LUBRICATION AND SCAVENGE OIL PUMP The lubrication and scavenge oil pump (LSOP) provides pressure for lubricating components, and a scavenge function to draw the oil back to the oil control module (OCM) and tank for reuse. The LSOP has seven stages. Six of the stages scavenge oil from the bearing compartments and gearboxes. The seventh stage pressurizes the oil, and sends it to the oil control manifold where it then goes directly to the main oil filter. Six magnetic chip collectors in each of the return lines catch metallic particles. The collectors catch ferrous metal particles in the scavenge oil, which can then be used to diagnose system problems. The chip collectors have quick removal type fittings for easy maintenance. Chip collectors are labeled according to the scavenge stage the oil is collected from: •
Front bearing compartment (FBC), including the fan drive gear system (FDGS) and bearings no. 1, no. 1.5, and no. 2
•
No. 3 bearing compartment (BC3)
•
No. 4 bearing compartment (BC4)
•
No. 5 and no.6 bearing compartments (BC5)
•
Main gearbox (MGB)
•
Angle gearbox (AGB)
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79-12
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
LEGEND Accessory Gear Box Bearing Compartment 3 Bearing Compartment 4 Bearing Compartment 5 Forward Bearing Compartment Main Gear Box
Housing
Chip Collector
CS1_CS3_7920_004
AGB BC3 BC4 BC5 FBC MGB
Figure 6: Lubrication and Scavenge Oil Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-13
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
OIL CONTROL MODULE The oil control module (OCM) is used to mount related oil system components. Various line replaceable units (LRUs) are centered around the body of the OCM to simplify maintenance. The OCM provides various oil passages between components. Oil flows into and through this module, using external oil tubes, and OCM housing passages.The oil metering plugs control the flow of oil and can be replaced to allow system pressures and engine flow to be adjusted if required. The following distribution system components are mounted on the OCM: •
Main oil filter (MOF)
•
Variable oil reduction valve (VORV)
•
Journal oil shuttle valve (JOSV)
•
Active oil damper valve (AODV)
•
Oil debris monitor (ODM)
•
Oil metering plugs
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79-14
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Active Oil Damper Valve (AODV)
Main Oil Filter
Oil Debris Monitor
Variable Oil Reduction Valve/Journal Oil Shuttle Valve (VORV/JOSV)
Oil Metering Plugs
CS1_CS3_7920_005
Oil Metering Plug
Figure 7: Oil Control Module Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-15
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
DEOILER During engine operation, sealing air flows into the bearing compartments. The sealing air must be vented to allow a continuous flow of air. The vented sealing air is referred to as breather air. The engine oil breather system removes air from the bearing compartments, separates breather air from the oil, and vents the air overboard. The deoiler receives air from the main gearbox, the no. 3 bearing compartment, and the deaerator in the oil tank. The deoiler separates the oil from the air. The oil returns to the gearbox for scavenging, and the air is vented overboard through the deoiler vent duct. The line replaceable deoiler drive oil seal creates a seal between the deoiler shaft and the main gearbox.
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79-16
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Deoiler Vent Duct
MAIN GEARBOX
CS1_CS3_7920_010
Deoiler
Figure 8: Deoiler Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-17
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
MAIN OIL FILTER The main oil filter assembly is a component of the OCM. It removes contaminants from the oil as it exits the LSOP. The main oil filter assembly consists of the main oil filter housing, machined into the OCM, a main oil filter element, and a main oil filter bypass valve. The main oil filter is covered by a thermal blanket for heat protection. The main oil filter element is a reverse flow type, meaning the oil flow path is from the inner diameter of the filter to the outside. If the main oil filter clogs, the oil is bypassed into a secondary filter within the main filter assembly. The main oil filter housing has two jackscrew lands to facilitate removal of the cover. The housing has a built-in drain and vent plugs to allow drainage before the cover is removed.
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79-18
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Vent Plug
Secondary Filter
Thermal Blanket Cover Oil Filter Cover
Drain Plug
Primary Filter Jackscrew Lands Thermal Blanket
CS1_CS3_7920_006
Oil Filter Bypass
Figure 9: Main Oil Filter Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-19
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
FAN DRIVE GEAR SYSTEM OIL PUMP The dual stage pump continuously draws oil from a dedicated auxiliary reservoir and compartment sump located in the front bearing compartment. The pump is installed in the no. 1 and no. 1.5 bearing support, and provides the journal bearings with an alternate oil supply during engine windmilling and negative-G conditions. It is turned by the fan drive shaft and rotates at fan speed.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-20
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
Auxiliary Reservoir
Fan Oil Pump
CS1_CS3_7920_012
No. 1 and No. 1.5 Bearing Support
Figure 10: Fan Drive Gear System Oil Pump Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-21
CS130_79_L2_Distribution_ComponentInformation.fm
79 - Oil 79-20 Distribution
DETAILED COMPONENT INFORMATION JOURNAL OIL SHUTTLE VALVE The journal oil shuttle valve (JOSV) is a mechanical, two position device that directs oil flow from the main oil or fan drive gear system oil supply to the journal bearings. The JOSV is controlled by comparing the main oil pressure against the gearbox vent pressure. When oil pressure is normal, the primary oil goes to the journal bearings. If oil pressure decreases, the JOSV sends oil from the fan drive gear system to the journal bearings. VARIABLE OIL REDUCTION VALVE The variable oil reduction valve (VORV) diverts the supply of oil from the FDGS to the oil tank. The valve is controlled by the electronic engine control (EEC) through an electrohydraulic servovalve (EHSV). The EEC receives valve position feedback from a linear variable differential transformer (LVDT). Maximum oil flow to lubricate the gear faces of the FDGS is required at takeoff only. At cruise, oil flow is reduced and sent back to the oil tank. Less oil flowing to the gears reduces the oil heat load and increases fan drive gearbox efficiency. When in the failsafe position the valve reverts to maximum oil flow.
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79-22
CS130_79_L3_Distribution_DetailedComponentInformation.fm
79 - Oil 79-20 Distribution
JOURNAL OIL SHUTTLE VALVE
From Windmill Pump
Gearbox Vent
From Main Oil Supply
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
JOURNAL OIL SHUTTLE VALVE
From Windmill Pump
LVDT
LVDT
T/M
T/M
Gearbox Vent
To Journal Bearings
ReturnTo Oil Tank
To Fan Drive Gear System
To Journal Bearings
ReturnTo Oil Tank
JOSV NORMAL FLOW
JOURNAL OIL SHUTTLE VALVE
Gearbox Vent
From Main Oil Supply JOURNAL OIL SHUTTLE VALVE
From Windmill Pump
From Main Oil Supply
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
LVDT
LVDT
T/M
T/M
Gearbox Vent
To Journal Bearings
ReturnTo Oil Tank
To Fan Drive Gear System VORV TAKEOFF
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
From Windmill Pump
From Main Oil Supply
From Fuel/ VARIABLE OIL Oil Heat REDUCTION Exchanger VALVE
To Fan Drive Gear System JOSV LOW MAIN OIL PRESSURE
To Journal Bearings
ReturnTo Oil Tank
To Fan Drive Gear System VORV CRUISE
LVDT T/M
Fan Drive Gear System Oil Pressure Return Vent Liner Variable Differential Transformer Torque Motor
NOTE JOSV shown in normal position. VORV shown in low flow position.
CS1_CS3_7921_001
LEGEND
Figure 11: Journal Oil Shuttle Valve and Variable Oil Reduction Valve Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-23
CS130_79_L3_Distribution_DetailedComponentInformation.fm
79 - Oil 79-20 Distribution
DETAILED DESCRIPTION The oil leaves the tank through an oil strainer, passes through an oil pressure element of the LSOP, then to the OCM where it is filtered by the main oil filter and distributed. If the filter begins clogging, an OIL FILTER IMPENDING BYPASS info message is shown. The filter bypasses at 55 psid as indicated by a L(R) OIL FILTER caution message and the secondary oil filter maintains a sufficient level of filtering. The no. 3 damper supply is fed just downstream of the filter and the remainder of the oil is then split between cooled and uncooled paths. The cooled path is delivered to the heat exchangers where it is cooled and returned to the OCM. From there, some of the cooled oil is supplied to the FDGS journal bearings, and the remainder of the cooled oil mixes with the uncooled oil and is distributed to lubricate the engine main shaft bearings and seals, FDGS gears, and both the accessory and angle gearboxes. After these components have been lubricated, the oil is scavenged by the LSOP and pumped back through the OCM to the oil tank, where it is deaerated. The LSOP has six scavenge pump elements that pull oil from the: •
Front bearing compartment servicing the FDGS and bearings no. 1, no. 1.5, and no. 2
•
No. 3 bearing compartment
•
No. 4 bearing compartment
•
No. 5 and no. 6 bearing compartment
•
Main gearbox
•
Angle gearbox
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
The oil drains into the tank, while the breather air/oil mixture leaves the tank and enters the gearbox. A rotating deoiler separates the air/oil mixture, allowing oil to be scavenged by the gearbox scavenge pump, and pass air to the overboard vent. Large particles could enter the oil supply beyond the main oil filter on the oil control module. Last chance oil strainers located in the supply lines prevent these particles from entering bearing compartments and clogging oil nozzles. The last chance strainers are located: •
FDGS gears
•
No. 1, no. 1.5, and no. 2 bearing compartments
•
No. 3 bearing compartment
•
No. 4 bearing compartment
•
No. 5 and no. 6 bearing compartments
•
Angle gear box
Heating and cooling of the engine fuel and oil is accomplished by modulating the flow of oil through a fuel/oil heat exchanger (FOHX) and air/oil heat exchanger (AOHX). The proportion of oil that flows through the two types of heat exchanger are controlled by the electronic engine control (EEC) and the fuel/oil heat exchanger bypass valve (FOHXBV). The FOHXBV is controlled based on the fuel temperature, ensuring that the engine oil and engine fuel temperatures are maintained within limits.
TECHNICAL TRAINING MANUAL
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79-24
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
OIL CONTROL MODULE VORV MOP
JOSV
No. 4
Journal Bearings
LUBRICATION AND SCAVENGE OIL PUMP
MOT
Pressure Valve De-aerator
Deoiler Overboard
A O P
MGB VFG
Oil Level Sensor
ODM
No. 3
No. 6
MOF
Active Oil Damper Valve
No. No. Gear Teeth No. 1 1.5 2 Auxiliary Dual-Stage Tank Windmill Pump
No. 5
OIL TANK
PMAG
FDGS AUXILIARY CIRCUIT LEGEND
ANGLE GEARBOX
FOHX FOHX MANIFOLD AIR/OIL HEAT EXCHANGER VFG AIR/OIL HEAT EXCHANGER
OIL/OIL HEAT EXCHANGER
FUEL/OIL MANIFOLD
MAIN GEARBOX
AOP FDGS FOHX JOSV MOF MOP MOT
Breather Air ODM VFG Scavenge PMAG VFG Oil Supply VFG Pressure Scavenge VORV Fan Drive Gear System Auxiliary Oil Pressure Fan Drive Gear System Fuel/Oil Heat Exchanger Journal Oil Shuttle Valve Main Oil Filter Main Oil Pressure Main Oil Temperature
Oil Debris Monitor Permanent Magnet Alternator Generator Variable Frequency Generator Variable Oil Reduction Valve Chip Collector Last Chance Screen Valve
CS1_CS3_7930_005
FOHX Bypass Valve
Figure 12: Distribution Schematic Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-25
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
OIL SYSTEM CONTROL
Variable Oil Reduction Valve
The EEC controls the following oil system components:
The VORV is an electromechanical valve that diverts oil from the FDGS to the oil tank. The EEC controls the dual-channel torque motor driven VORV, while a linear variable differential transformer (LVDT) provides position feedback. Oil flow to lubricate the gear faces is only required at takeoff. At cruise, less oil flow is required, so the VORV diverts some of the oil back to the oil tank, to maximize efficiency during cruise conditions. The amount of oil required is a function of the load placed on the FDGS. The amount of oil is scheduled as a function of horsepower extracted by the fan.
Active Oil Damper Valve The no. 3 bearing active oil damper valve (AODV) controls the oil supply to the no. 3 bearing damper. The oil supply is controlled to: •
Optimize no.3 bearing loads during all phases of operation
•
Limit high-pressure spool vibration
•
Provide bowed rotor protection at sub-idle operation
The solenoid operated valve is controlled by either channel of the EEC based on N2 rpm. The valve is open during start up to N2 idle speed (62%). The valve is energized closed up to 85% N2, then reopens. Fuel/Oil Heat Exchanger Bypass Valve The FOHXBV is an electromechanical modulating valve that controls the amount of oil going to the FOHX and/or the AOHX.
The fan thrust is calculated based on shaft speed, air pressure, and temperature entering the fan. Due to the similarity in the thrust and P3 burner pressure sensor characteristic response over the flight envelope, the VORV is scheduled as a function of P3 for above idle operation. If the VORV is stuck in the low flow setting, then inadequate lubrication at high engine power settings may cause degradation of the FDGS gears. If stuck in the high flow setting, the oil temperature reads high at low power settings.
The EEC receives the fuel temperature sensor signal and varies the amount of oil going to each heat exchanger, based on the fuel temperature sensor readings. The FOHXBV supplies a minimum of 25%, and a maximum of approximately 100% of the oil flow to the AOHX. Under normal operating conditions, the FOHXBV splits oil flow as follows: •
75% is supplied to the fuel/oil heat exchanger
•
25% is suppled to the air/oil heat exchanger
The FOHXBV increases oil flow through the FOHX if the fuel temperature is too low. The increased oil flow through the FOHX raises the fuel temperature.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
From VFG EEC
To VFG Air/Oil Heat Exchanger
VFG OIL/OIL HEAT EXCHANGER
AIR/OIL HEAT EXCHANGER
Channel A
FOHXBV T/M
OCM
To Oil System
LVDT
Channel B
To No. 3 Bearing
AODV
To FDGS
Fuel FUEL/OIL HEAT EXCHANGER
To Oil Tank
FUEL TEMPERATURE SENSOR
T/M
VORV
From Oil Pump
P3 BURNER PRESSURE SENSOR
LEGEND
CS1_CS3_7930_006
AODV EEC FDGS FOHXBV LVDT OCM T/M VFG VORV
Fuel Oil VFG Oil Active Oil Damper Valve Electronic Engine Control Fan Drive Gear System Fuel/Oil Heat Exchanger Bypass Valve Linear Variable Differential Transformer Oil Control Module Torque Motor Variable Frequency Generator Variable Oil Reduction Valve CAN BUS
Figure 13: Oil System Control Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-27
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
Fan Drive Gear System Oil Circuit The fan drive gear system (FDGS) journal bearings are lubricated with pressurized oil from a manifold at the rear of the FDGS. The FDGS oil supply is provided by the variable oil reduction valve (VORV) and the journal oil shuttle valve (JOSV). Both valves ensure the fan drive gear system (FDGS) receives the correct amount of oil when required. During normal operating conditions, the main oil pump pressurizes the oil system through the VORV and the JOSV returns the sump oil to the oil tank. A dual-stage fan-driven pump acts as a supplemental scavenge pump, drawing oil from the sump and returning it to the oil tank. During engine windmilling in flight, on the ground, or during negative-G conditions, the main oil pressure decreases. The journal oil shuttle valve moves to isolate the journal bearing oil flow from the main oil supply, allowing the fan-driven oil pump to supply oil from the sump to the journal bearings. In a negative-G condition, the fan-driven pump draws oil from the secondary sump to supply oil to the journal bearings. The operation of the journal oil shuttle valve and fan-driven oil pump are verified on every flight via the auxiliary oil pressure sensor. This detects both positive oil flow from the fan-driven oil pump, and shuttle valve movement from the main oil pump supply position to the fan-driven pump oil supply position. Rotation of the fan for maintenance does not impact the oil system components. A failure of the fan-driven oil pump or JOSV is detected by the auxiliary oil pressure sensor and indicated at the next engine start.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-20 Distribution
NORMAL OPERATION Journal Bearing Oil Filter
NEGATIVE-G OR WINDMILL OPERATION Journal Bearing Oil Filter
VORV/JOSV
NO. 1/1.5 BEARING SUPPORT
VORV/JOSV
NO. 1/1.5 BEARING SUPPORT
Oil Manifold
Oil Manifold OIL TANK
OIL TANK
FDGS/ Journal Bearings
FDGS/ Journal Bearings
Gutter
Gutter
Lubrication Pump
Secondary Sump
Scavenge Pump
Fan Oil Pump
Secondary Sump
Lubrication Pump
Scavenge Pump
Fan Oil Pump
LEGEND
FDGS VORV/JOSV
Sump
FDGS Pressure Scavenge Supply Fan Drive Gear System Variable Oil Reduction Valve/Journal Oil Shuttle Valve Auxiliary Oil Pressure Sensor
CS1_CS3_7930_015
Sump
Figure 14: Fan Drive Gear System Operation Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-29
CS130_79_L3_Distribution_DetailedDescription.fm
79 - Oil 79-30 Indicating
79-30 INDICATING GENERAL DESCRIPTION The oil indicating system provides oil temperature, pressure, and quantity information for display in the flight deck. All of the sensors and probes are line replaceable units (LRUs). The oil quantity sensor is mounted on top of the oil tank. The oil quantity sensor provides the quantity of engine oil in the oil tank. The dual-channel oil pressure sensor measures the pressure output of the lubrication and scavenge oil pump (LSOP). The dual-channel oil temperature probe measures the temperature of the scavenge oil returning to the oil tank. The dual-channel oil filter differential pressure sensor monitors the differential pressure across the main oil filter, and provides an indication of an impending filter bypass. The oil debris monitor (ODM) detects both ferrous, and non-ferrous debris in the oil system. The presence of metallic particles is detected by a sense coil in the ODM. This produces a signal recorded by the prognostics and health management unit (PHMU), which classifies the particles according to size, and whether it is ferrous or non-ferrous. The PHMU records the data for engine health monitoring purposes. A dual-element auxiliary oil pressure sensor monitors oil pressure in the fan drive gear system (FDGS). The oil circuit detects failures of the dual-stage pump and the journal oil shuttle valve (JOSV). The sensors and probes supply information to the electronic engine control (EEC). The EEC provides the information for display on the engine indication and crew alerting system (EICAS), the STATUS synoptic page, the onboard maintenance system (OMS), and the health management unit (HMU).
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CS130_79_L2_Indicating_General.fm
79 - Oil 79-30 Indicating
Oil Quantity Sensor
TO 90.8
90.8
18.6
18.6
N1
516 OIL TANK
EGT ELECTRONIC ENGINE CONTROL
Oil Filter Differential Pressure Sensor
522
62.1 220 103 113
Oil Pressure Sensor
N2 FF (KPH) OIL TEMP OIL PRESS
62.0 220 103 113
EICAS Oil Temperature Sensor
OIL CONTROL MODULE
Oil Debris Monitor
Excitation
ENGINE
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
103 113 21.8
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (L)
103 113 21.8
SYNOPTIC PAGE – STATUS LEGEND ARINC 429 CAN BUS Analog
OMS HMU
CS1_CS3_7930_016
AUXILIARY OIL PRESSURE SENSOR
Figure 15: Oil Indicating Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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CS130_79_L2_Indicating_General.fm
79 - Oil 79-30 Indicating
COMPONENT LOCATION The following oil indicating component is located on the oil tank: •
Oil quantity sensor
The following components are located on the oil control module (OCM): •
Oil filter differential pressure sensor
•
Oil debris monitor
•
Oil pressure sensor
•
Oil temperature sensor
The following auxiliary oil system component is located at the 5 o’clock position on the fan intermediate case. •
Auxiliary oil pressure sensor
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79 - Oil 79-30 Indicating
Oil Quantity Sensor Oil Filter Differential Pressure Sensor
A
Oil Pressure Sensor
B D
Oil Temperature Sensor A
OIL TANK
B
OIL CONTROL MODULE
C
OIL DEBRIS MONITOR
D
AUXILIARY OIL PRESSURE SENSOR
CS1_CS3_7930_008
C
Figure 16: Oil System Indicating Components Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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CS130_79_L2_Indicating_ComponentLocation.fm
79 - Oil 79-30 Indicating
DETAILED COMPONENT INFORMATION OIL DEBRIS MONITOR The ODM sensor core consists of a magnetic coil assembly with two field coils and one sense coil. The coil assembly receives an excitation signal from the PHMU. Two equal and opposing magnetic fields are generated, which balance in the center. Passage of a metallic particle disrupts the balance of the fields seen by the sense coil, and results in a characteristic signal. The signal resembles a sine wave: •
Amplitude is proportional to particle size
•
Phase of ferrous particle signal versus non-ferrous particle signal is offset by 90°
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79 - Oil 79-30 Indicating
Field Coils
Sense Coil
Coil Assembly
Particle
Signal (volt)
Signal ~ Particle Size
LEGEND Ferrous Non-Ferrous
Excitation Signal Oil Passage Tube
Coils
OIL DEBRIS MONITOR CROSS – SECTION VIEW
CS1_CS3_7930_002
PROGNOSTICS AND HEALTH MANAGEMENT UNIT
Figure 17: Oil Debris Monitor Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79 - Oil 79-30 Indicating
CONTROLS AND INDICATIONS EICAS The oil temperature and pressure are displayed on the EICAS page. STATUS The oil temperature, pressure, and quantity are displayed on the STATUS synoptic page.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79 - Oil 79-30 Indicating
TO 90.8
90.8
18.6
18.6
OIL TEMPERATURE Symbol
N1
516
Description
103
Oil temperature normal range
158
Oil temperature above high oil temperature yellow line
183
Oil temperature above high oil temperature red line threshold
35
522
èèè
Oil temperature below low oil temperature threshold Oil temperature invalid
EGT N2 FF (KPH) OIL TEMP OIL PRESS
62.0 220 103 113
OIL PRESSURE Symbol
81
AIR IR
DOOR OR
ELEC C
FLT CTR CTRL TRL L
FUEL FUE EL
HYD YD
AVIONIC IONIC
INFO
CB
TAT SAT
Oil pressure above high oil pressure threshold
20
Oil pressure below low oil pressure threshold
èèè
Oil pressure invalid
15°C 15°C
OIL QUANTITY Symbol
ENGINE
103 113 21.8
Oil pressure normal range
203
EICAS
STATUS
Description
OIL TEMP (°C) OIL PRESS (PSI) OIL QTY (QTS)
19.2
103 113 21.8
Description Normal
2.0
Below threshold
èèè
Invalid
STATUS PAGE
CS1_CS3_7930_018
62.1 220 103 113
Figure 18: Oil Indications Copyright © Bombardier Inc. CS130-21.07-06-00-121418
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79-37
CS130_79_L2_Indicating_ControlsIndications.fm
79 - Oil 79-30 Indicating
DETAILED DESCRIPTION The oil quantity sensor sends an oil quantity signal to channel B of the EEC. When the main oil tank quantity is at or below 9.45 L (10 qt), a cyan L (R) ENG OIL LO QTY message is displayed on the EICAS page. The oil filter differential pressure sensor signal is sent to both channels of the EEC. When the differential pressure across the filter is more than 35 psid at rated flow conditions, an OIL FILTER IMPENDING BYPASS INFO message is displayed. When the differential pressure across the oil filter element is more than 55 psid, the filter bypass valve opens and an L (R) ENG OIL FILTER caution message is displayed.
An AUX OIL PRESS MON INOP INFO message is displayed when the auxiliary oil pressure sensor fails. Table 1: Oil System Limits OIL SYSTEM LIMITS PARAMETER Oil Quantity Oil Filter Bypass
LIMIT
WARNING
< 9.45 L (10 qt)
L(R) ENG OIL LO QTY advisory message.
35 psid
Oil filter impending bypass INFO message.
55 psid
L(R) ENG OIL FILTER caution message.
The oil pressure sensor sends signals to both channels of the EEC. The oil temperature probe sends oil temperature information to both channels of the EEC.
Oil Temperature
An ENG OIL LO TEMP caution message is displayed if the throttles are advanced beyond 50% N1 before the oil temperature has reached 49°C (120°F).
L(R) ENG EXCEEDANCE caution message.
Minimum oil temperature for starting is - 40° C (40° F). Operation at maximum oil temperature of 174° C (345° F) cannot exceed 20 minutes. Oil Pressure
The operation of the JOSV and the fan-driven oil pump are verified by the auxiliary oil pressure sensor, which detects positive oil flow from the fan-driven oil pump. It also detects shuttle valve movement from the main oil pump supply position to the fan-driven pump oil supply position.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
ENG OIL LO TEMP caution message if throttles are advanced 50%.
> 152°C (306°F)
If the oil temperature exceeds 152°C, a L (R) ENG EXCEEDENCE caution message is displayed. The ODM receives an excitation signal from the PHMU. The ODM detects ferrous and non-ferrous debris in the oil system, and sends a signal back to the PHMU. The PHMU supplies the ODM information to both channels of the EEC. If the amount of debris exceeds a predefined limit, an OIL DEBRIS ABOVE LIMIT INFO message is displayed.
< 49°C (120°F)
TECHNICAL TRAINING MANUAL
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Operation at maximum oil pressure cannot exceed 10 minutes.
ENG OIL LO PRESS warning message when oil is below minimum threshold with the engine running. CS1_CS3_TB79_002
79-38
CS130_79_L3_Indicating_DetailedDescription.fm
79 - Oil 79-30 Indicating
TO OIL TEMPERATURE PROBE
91.7
91.7
21.2
21.2
ELECTRONIC ENGINE CONTROL
N1
618 OIL PRESSURE SENSOR
OIL FILTER DIFFERENTIAL PRESSURE SENSOR
Channel A
624
ENG OIL LO TEMP EMER LTS OFF EMER DEPRESS ON HYD PUMP 2B ON APU ON CABIN PRESS MAN PARK BRAKE ON
EGT
68.2 280 72 85
N2 FF (PPH) OIL TEMP OIL PRESS
68.3 285 48 98
GEAR
DN DN DN
Channel B
IN
TOTAL FUEL (KG) 2490 SYN
OIL QUANTITY SENSOR
0
5070
SLAT / FLAP
0
2580 EICAS
OMS Excitation PROGNOSTICS AND HEALTH MANAGEMENT UNIT
HMU
OIL DEBRIS MONITOR AUXILIARY OIL PRESSURE SENSOR CS1_CS3_7930_017
LEGEND ARINC 429 CAN BUS Analog
Figure 19: Oil Indicating Detailed Description Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-39
CS130_79_L3_Indicating_DetailedDescription.fm
79 - Oil 79-30 Indicating
OIL PRESSURE WARNING LIMITS The EEC sends a signal to EICAS for engine oil pressure. A red L (R) ENG OIL LO PRESS EICAS warning message is shown if the oil pressure drops below the minimum threshold while the engine is running. During an oil pressure test, the oil pressure should fall within the trim band range. If the oil pressure falls between the trim band and the high oil pressure or low oil pressure yellow limit, the oil metering plugs on the oil control module (OCM) can be replaced with different sized metering plugs to bring the oil pressure into the trim band range. This can only be done after consultation with Pratt & Whitney.
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79 - Oil 79-30 Indicating
MAIN OIL PRESSURE - MOP (PSIG)
L ENG EXCEEDANCE 200
n
eratio
al Op
orm um N
Maxim 150
tion
al Opera
m Norm Minimu 100
m Limit
Minimu 50
L ENG OIL PRESS 0 13000 (53)
14000 (57)
15000 (61)
16000 (65)
17000 (69)
18000 (74)
19000 (78)
20000 (82)
21000 (86)
22000 (90)
23000 (94)
24000 (98)
CS1_CS3_7930_020
HIGH ROTOR SPEED – N2 (RPM/1000) (% RPM)
Figure 20: Oil Pressure Warning Limits Copyright © Bombardier Inc. CS130-21.07-06-00-121418
TECHNICAL TRAINING MANUAL
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79-41
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79 - Oil 79-30 Indicating
MONITORING AND TESTS The following page provides crew alerting system (CAS) and INFO messages for the oil indicating system.
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79 - Oil 79-30 Indicating
CAS MESSAGES Table 2: WARNING Messages MESSAGE
LOGIC
L ENG OIL PRESS
Left engine oil pressure is below or above the normal range.
R ENG OIL PRESS
Right engine oil pressure is below or above the normal range.
Table 3: CAUTION Messages MESSAGE
LOGIC
ENG OIL LO TEMP
Either engine oil too cold to allow a high thrust to be set.
L ENG OIL FILTER
Left engine oil filter is bypassed (clogged filter).
R ENG OIL FILTER
Right engine oil filter is bypassed (clogged filter).
L ENG EXCEEDANCE
L N1 or L N2 or L EGT or L oil temperature is above the threshold.
L ENG EXCEEDANCE
R N1 or R N2 or R EGT or R oil temperature is above the threshold.
Table 4: ADVISORY Messages MESSAGE
LOGIC
L ENG OIL LO QTY
Left engine oil, low quantity detected.
R ENG OIL LO QTY
Right engine oil, low quantity detected.
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79 - Oil 79-30 Indicating
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79 - Oil 79-30 Indicating
Table 5: INFO Messages Table 5: INFO Messages MESSAGE 79 L ENGINE FAULT - OIL DEBRIS ABOVE LIMIT
MESSAGE
LOGIC INFO message will be set when an oil debris level above limit has been detected in either channel A or B by the left engine FADEC system.
79 R ENGINE FAULT - OIL DEBRIS ABOVE LIMIT
INFO message will be set when an oil debris level above limit has been detected in either channel A or B by the right engine FADEC system.
79 L ENGINE FAULT- OIL DEBRIS MON INOP
FADEC logic has determined that the sensor has failed.
79 R ENGINE FAULT- OIL DEBRIS MON INOP
FADEC logic has determined that the sensor has failed.
79 L ENGINE FAULT - OIL QTY FADEC logic has determined that the oil SNSR INOP quantity signal from channel A and/or channel B has failed.
79 L ENGINE FAULT - VORV OPER DEGRADED
FADEC logic has determined that the oil filter P signal in channel A and/or channel B has failed.
79 R ENGINE FAULT - OIL FILTER SENSOR INOP
FADEC logic has determined that the oil filter P signal in channel A and/or channel B has failed.
79 L ENGINE FAULT - OIL FILTER IMPENDING BYPASS
INFO message will be set by the left engine when the oil filter impending bypass is detected but there is no actual bypass.
79 R ENGINE FAULT - OIL FILTER IMPENDING BYPASS
INFO message will be set by the right engine when the oil filter impending bypass is detected but there is no actual bypass.
Copyright © Bombardier Inc. CS130-21.07-06-00-121418
INFO message will be set by the left engine in any of the following conditions: - Any channel of the VORV feedback sensor is failed - The active EEC channel has detected a cross-check failure on the VORV feedback sensor
79 R ENGINE FAULT - VORV OPER DEGRADED
INFO message will be set by the right engine in any of the following conditions: - Any channel of the VORV feedback sensor is failed - The active EEC channel has detected a cross-check failure on the VORV feedback sensor
79 L ENGINE FAULT - BRG DAMPER VLV INOP
INFO message will be set by the left engine in any of the following conditions: - Both EEC channels are unable to command bearing damper valve solenoid
79 R ENGINE FAULT - OIL QTY FADEC logic has determined that the oil SNSR INOP quantity signal from channel A and/or channel B has failed. 79 L ENGINE FAULT - OIL FILTER SENSOR INOP
LOGIC
- Any channel control processor failed and remaining channel unable to command bearing damper valve solenoid 79 R ENGINE FAULT - BRG DAMPER VLV INOP
TECHNICAL TRAINING MANUAL
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INFO message will be set by the right engine in any of the following conditions: - Both EEC channels are unable to command bearing damper valve solenoid - Any channel control processor failed and remaining channel unable to command bearing damper valve solenoid
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79 - Oil 79-30 Indicating
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79 - Oil 79-30 Indicating
Table 5: INFO Messages MESSAGE
LOGIC
79 L ENGINE FAULT - AUX OIL An INFO message is set by the left engine PRESS MON INOP when the auxiliary oil pressure sensor fails. This INFO message is masked if the PHMU fails. 79 R ENGINE FAULT - AUX OIL An INFO message is set by the right engine PRESS MON INOP when the auxiliary oil pressure sensor fails. This INFO message is masked if the PHMU fails.
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CS130_79_ILB.fm
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