Dash 8 Series Maintenance Training Manual Vol1 Part1: Revision: 5.1

DASH 8 SERIES MAINTENANCE TRAINING MANUAL VOL1 PART1 Revision: 5.1 FlightSafety International,Inc. Toronto Learning Centre 95 Garratt Boulevard Downsview, Ontario, Canada (416) 638-9313 www.FlightSafety.com

NOTICE The material contained in this training manual is based on information obtained from the aircraft manufacturer’s pilot manuals and maintenance manuals. It is to be used for familiarization and training purposes only. At the time of printing it contained then-current information. In the event of conflict between data provided herein and that in publications issued by the manufacturer, that of the manufacturer shall take precedence. We at FlightSafety want you to have the best training possible. We welcome any suggestions you might have for improving this manual or any other aspect of our training program.

FlightSafety international

INSTRUCTIONAL SYSTEM DIVISION

8900 Trinity Blvd. Hurst, Texas 76053

(817) 595-5450

DASH 8 MAINTENANCE TRAINING MANUAL, VOLUME 1 Record of Revision No. 5.1 This is Revision No. 5.1 to the Dash 8 Maintenance Training Manual, Volume 1. Pages are dated Revision 5 in the lower left or right hand corner. Each page updates the previous version. The portion of the text or figure affected by the current revision is indicated by a solid vertical line in the margin. A vertical line adjacent to blank space means that material has been deleted. The changes made in this revision will be further explained at the appropriate time in the training course.

safety begins with a well-trained maintenance technician...

CONTENTS VOLUME 1 CHAPTER TITLE

ATA NUMBER

INTRODUCTION

1

ATA100

2

AIRCRAFT GENERAL

4

AIR CONDITIONING (SERIES 100/100A)

21

AIR CONDITIONING (SERIES 300A)

21A

AIR CONDITIONING (SERIES 200)

21B

AVIONICS

22

ELECTRICAL POWER

24

FIRE PROTECTION

26

FLIGHT CONTROLS

27

FUEL

28

HYDRAULIC

29

ICE AND RAIN PROTECTION

30

LIGHTING

31

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INTRODUCTION

This training manual provides a description of the major airframe and engine systems installed in the Dash 8. This information is intended as an instructional aid only; it does not supersede, nor is it meant to substitute for, any of the manufacturer’s maintenance or operating manuals. This material has been prepared from the basic design data, and all subsequent changes in airplane appearance or system operation will be covered during academic training and subsequent revisions to this manual.

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T h i s m a n u a l i s i n t wo p a r t s : t h e b a s i c Maintenance Training Manual and the supplemental Maintenace Schematic Manual (MSM). The MSM contains schematics to be used only as a tool in understanding a system. They are not kept current; the corresponding schematics(s) in the manufacturers’s Maintenance Manual must be used when performing maintenance. The basic text contains references to schematics found in the MSM. MSM schematics are designated by an “S” at the end of the figure n u m b e r. A s a n e x a m p l e , t h e fi r s t M S M schematic pertaining to Chapter 29 is Figure 29-1S. Succeeding figures are -2S, -3S, etc. Also, as an aid to finding the corresponding schematic in the manufacturer’s Maintenance Manual, the title of each MSM figure contains the Maintenance Manual chapter/section number and figure number in parentheses; for example, the title of Figure 29-1S is “No. 1 Hydraulic Power System Schematic (29-10-00, Figure 1).” The second chapter of this manual, “ATA 100,” is an introduction to the Air Transport Association format for aircraft maintenance manuals. It is intended to describe simply the basic format for all ATA 100 Maintenance Manual chapters and also to explain where variation may exist from one manufacturer to another. Each chapter following “ATA 100” of this book has listed on the divider tab the ATA chapter(s) included, such as “24 Electrical.” In some cases it was appropriate, for training purposes, to include more than one ATA chapter in one chapter of this book, such as Chapters 4 through 12 in “Aircraft General.” The tab marked “4-12 Aircraft General.” indicates that applicable ATA 100 Maintenance Manual Chapters 4 through 12 are covered in that chapter. Any chapter not included the manufacturer’s Maintenance Manual for that particular airplane is not included in that chapter of this training manual. Appendix A in this manual shows the circuitbreaker panels. Appendix B displays all light indications and should be folded out for reference while reading this manual. Appendix C contains a pictorial Walkaround on a Dash 8. 1-2

The goal of this course is to provide the very best training possible for the clients in our maintenance initial program. So that there is no uncertainty about what is expected of the client, the following basic objectives are presented for this course. Given the Maintenance Manual , class notes, and this training manual (as specified by the FlightSafety instructor), the client will be able to pass a written examination upon completion of this course to the grading level prescribed by the FlightSafety Director of Training. The maintenance technician will be able to: • Outline the ATA 100 system of maintenance documentation, including the major chapter headings and symbology. • Describe the meaning and application of each piece of manufacturer’s maintenance documentation and use the documentation in practical applications, • Outline the manufacturer’s spares, technical support, and warranty procedures, and name the principal contacts. • Outline the recommended maintenance schedule and the appplicable options. • Locate major components without references to documentation and other components with the aid of documentation. • Describe the operation of all major systems in the normal and various abnormal operating modes. • Perform maintenance preflight and postflight inspections. • Perform selected normal and emergency cockpit procedures as required for engine start/run-up, APU start, battery check, airplane taxiing, etc. (requires use of a simulator). The Flight Safety instructor will modify the stated overall objective conditions and criteria to satisfy selected performance requirements, when appropriate. The performance levels specified will not vary from those directed by the FlightSafety Director of Training.

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A glossary of abbreviated terms found in this manual is shown on the following pages. REFERENCE

DESCRIPTION

LOCATION

ACM

Air Cycle Machine

Air Conditioning/Pressurization

ADC

Air Data Computer

Flight Instruments

ADI

Attitude Director Indicator

Flight Instruments

AFCS

Automatic Flight Control System

Auto Flight

APU

Auxiliary Power Unit

Auxiliary Power

ASCB

Avionics Standard Communications Bus (Serial)

Autopilot Flight Instruments

BBPU

Bus Bar Protection Unit

Electrical (AC and DC)

BITE

Built-In Test Equipment

Electrical/APU

CLA

Condition Lever Angle (Pitch)

Propeller/Powerplant

CPU

Central Processor Unit

Auto Flight

CVR

Cockpit Voice Recorder

Avionics/Communications

DFDR

Digital Flight Data Recorder

Avionics/Communications

EADI

Electronic Attitude Director Indicator

Flight Instruments

ECU

Electronic Control Unit

Engine

EFIS

Electronic Flight Instrument System

Flight Instruments

EHSI

Electronic Horizontal Situation Indicator

Flight Instruments

ELT

Emergency Locator Transmitter

Avionics/Communications

ESU

Electronic Sequence Unit

Auxiliary Power

FDAU

Flight Data Acquisition Unit

Avionics/Communications

FGC

Flight Guidance Computer

Automatic Flight

FPU

Flap Power Unit

Flight Controls

GCU

Generator Control Unit

Powerplant Electrical (AC)

GPWS

Ground Proximity Warning System

Avionics/Communications

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HMU

Hydromechanical Fuel Control Unit

Powerplants

HSI

Horizontal Situation Indicator

Flight Instruments

IMU

Inertial Measurement Unit

Flight Instruments

IVSI

Inertial Vertical Speed Indicator (Instantaneous)

Flight Instruments

PCU

Propeller Control Unit

Powerplants/Each Nacelle

PFCS

Powered Flight Control Surface Indicator

Flight Instrument

PLA

Power Lever Angle

Powerplants

PSEU

Proximity Switch Electronic Unit

Doors—Flight-Taxi Switch Flaps—Ground Spoilers Gear—Power Levers

PSU

Passenger Service Unit

Aircraft General

PTT

Press-to-Talk Switch

Avionics/Communications

PTU

Hydraulic Power Transfer Unit

Hydraulic Power/Landing Gear

RMI

Radio Magnetic Indicator

Flight Instruments

SCU

Signal Conditioner Unit

Powerplant/Each Nacelle

SPU

Standby Power System

Hydraulic Power

TRU

Transformer-Rectifier Unit

Electrical

ULD

Underwater Locating Device

Avionics/Communications

WOW

Weight on Wheels

XFMR

Transformer

Electrical

TCS

Touch Control Steering

Automatic Flight

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ATA 100 CONTENTS Page INTRODUCTION ................................................................................................................... 2-1 GENERAL............................................................................................................................... 2-1 DOCUMENTATION............................................................................................................... 2-2 General............................................................................................................................. 2-2 Revisions.......................................................................................................................... 2-2 Customized Chapters ....................................................................................................... 2-2 Associated Manuals ......................................................................................................... 2-3 ATA Codes ....................................................................................................................... 2-3 Updated MAS .................................................................................................................. 2-6 HOW TO USE THE MAINTENANCE MANUAL ................................................................... 2-7 Division of Subject Matter............................................................................................... 2-7 Standard Numbering System ........................................................................................... 2-7 List of Effective Pages ..................................................................................................... 2-8 SUMMARY............................................................................................................................. 2-8 DASH 8 SCHEDULED MAINTENANCE PROGRAM...................................................... 2-10 Systems.......................................................................................................................... 2-10 Structures ....................................................................................................................... 2-10 Airworthiness Limitations ............................................................................................. 2-10 Related Documentation.................................................................................................. 2-11 Implications of Repairs to Damage Tolerant Structure.................................................. 2-11

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ILLUSTRATION Figure 2-1

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Title

Page

ATA 100 Numbering ................................................................................................ 2-9

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ATA 100

INTRODUCTION The purpose of this chapter is to describe the arrangement, numbering system, and special features of the Air Transport Association format for aircraft maintenance manuals. To take advantage of all the material presented in an ATA 100-format manual, the maintenance technician must become thoroughly familiar with the outline and contents presented for any given airplane.

GENERAL ATA Specification No. 100 is issued by the Air Transport Association of America as the Specification for Manufacturers’ Technical Data. It establishes a standard for the presentation of certain data produced by aircraft, engine, and component manufacturers required for the support of their respective products. Under this format, the maintenance manual is broken down into standard chapters as defined

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by ATA 100. Each chapter covers a specific area of maintenance information, such as Chapter 10, “Parking and Mooring,” or a specific system, such as Chapter 32, “Landing Gear.” All data pertaining to a given system is located within its chapter, regardless of whether it is mechanical, hydraulic, or electrical in nature. The chapters are usually arranged in alpha betical order.

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DOCUMENTATION GENERAL The Dash 8 Maintenance Manual is prepared from manufacturer’s technical data in accordance with the Air Transport Association. The Maintenance Manual provides sufficient information to enable a mechanic who is unfamiliar with the airplane to service, troubleshoot, test, adjust, and repair systems and to remove and install any unit normally requiring such action on the line or in the maintenance hangar.

update the Dash 8 Maintenance Manual. Additions, deletions, or revisions to the text are identified in the Table of Contents and on the text page by a black bar in the left margin of the page.

Temporary Revisions Temporary Revisions are printed on yellow paper and notify operators rapidly of important changes or provide advance information on some equipment or modifications. The Temporary Revisions are filed in the manual as instructed in the Temporary Revision.

Record of Revisions REVISIONS General ATA 100 allows the manufacturers a great deal of leeway or freedom in the area of Maintenance Manual revisions and their disseminations. Virtually every aircraft manufacturer has a system different from any other manufacturer; some differences are great while others are barely noticeable, but all are intended to get maintenance information, routine or vital, to the field in a timely manner. Because changes, particularly new temporary changes, may be vital to ground and/or airborne safety, the maintenance technician should be thoroughly familiar with the methodology used by a particular manufacturer to incorporate changes into the Maintenance Manual. The manufacturer’s methods are listed in detail in the Maintenance Manual introduction for a given airplane. Two types of revisions are issued for the Dash 8–normal and temporary.

The Record of Revisions provides a place for the responsible individual to record the successive revision numbers, dates inserted, and his initials against the appropriate revision number. If the revision is inserted by the factory for a reprint of the manual, the revision record will show the revisions already incorporated.

Temporary Revision Index Temporary Revisions are recorded on the Temporary Revision Index. If the revision was incorporated by the manufacturer, it will be annotated as such.

CUSTOMIZED CHAPTERS Dash 8 Maintenance Manual chapters covering avionics (Chapters 22, 23, and 34) and furnishings (Chapter 25) are contained in a separate Customized Chapters Volume (CCV). The content of the CCV varies with each operator to reflect the avionics and furnishings configurations of his particular airplane.

Normal Revisions Normal Revisions are printed on white paper and issued to qualified holders as required to

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ASSOCIATED MANUALS

• Minimum Equipment List (DOT and FAA) ..................... PSM 1-8-16

Associated manuals for reference on Dash 8 maintenance include:

• MEL Procedures Manual (DOT and FAA)................... PSM 1-8-16A

• Aircraft Operating Data .......... PSM 1-8-1

• PW 120 Engine Maintenance Manual

• Flight Manual....................... PSM 1-8-1A

• 14SF-7 Propeller Maintenance Manual

• Maintenance Manual............... PSM 1-8-2

• Dash 8 Operators' Conferences

• DHC-8 Engine Rigging Manual.................. PSM 1-8-2ER • Ramp Servicing Manual........ PSM 1-8-2S • Tools and Equipment Manual .................................. PSM 1-8-2T

• March 25-27, 1986 and subsequent • Reliability and Maintenance Ability Statistical Analysis Dash 8 G.2 • Maintenance Tips • Service Review Operators’ Digest

• Wiring Diagrams Manual................................. PSM 1-8-2W

ATA CODES

• Structural Rigging Manual .................................... PSM 1-8-3

05 MAINTENANCE CHECKS

• Illustrated Parts Manual ......... PSM 1-8-4 • Component Maintenance Manual .................................. PSM 1-8-6* • Repair and Overhaul Agencies Manual................................ PSM 1-8-6A* • Maintenance Program Manual .................................... PSM 1-8-7 • Non-Destructive Testing Manual.................................. PSM 1-8-7A • Equalized Maintenance Program Manual .................................. PSM 1-8-7E • Weight and Balance Manual .................................... PSM 1-8-8 • Cargo Loading Manual.................................. PSM 1-8-8A • Power Plant Build-up Manual .................................. PSM 1-8-10 • Modifications and Options Manual .................................. PSM 1-8-12 • Crash–Fire–Rescue Information Manual .................................. PSM 1-8-14 • Vendor Warranties Manual .................................. PSM 1-8-15 * Available for component repair and overhaul

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0551 0552 0554 0555

LIGHTNING STRIKE BIRD STRIKE TURBULENCE HARD/OVERWEIGHT LANDING

21 AIR CONDITIONING 2110 2120 2130 2140 2150 2160 2170

COMPRESSION DISTRIBUTION (Eye Balls/Noises) PRESSURIZATION CONTROL HEATING COOLING (Packs—ACM) TEMPERATURE CONTROL WATER SEPARATOR

22 AUTO FLIGHT 2210 AUTOPILOT/YAW DAMPER 2220 SPEED–ATTITUDE CORRECTION (Mach Trim) 2230 AUTO THROTTLE 2240 SYSTEM MONITOR 2250 AERODYNAMIC LOAD ALLEVIATING (Active Controls)

23 COMMUNICATIONS 2310 SPEECH COMMUNICATION GENERAL 2311 HIGH FREQUENCY (HF) 2312 VERY HIGH FREQUENCY (VHF) 2-3

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2313 TELEPHONES (AIRFONE) 2320 DATA TRANSMISSION AND AUTO CALLING 2321 SELCAL 2322 DATA LINK 2330 PASSENGER ADDRESS AND ENTERTAINMENT 2331 PASSENGER ADDRESS 2332 MOVIES 2333 AUDIO ENTERTAINMENT 2340 INTERPHONE (Cabin, Service) 2350 AUDIO INTEGRATING (Fit Interphone) 2360 STATIC DISCHARGING (Wicks) 2370 AUDIO AND VIDEO MONITORING (Voice Recorder) 2380 INTEGRATED AUTOMATIC TUNING

24 ELECTRICAL POWER 2410 2421 2422 2430 2440 2450

GENERATOR DRIVE (CSD/IDG) AC GENERATOR SYSTEM AC INVERTER SYSTEM DC GENERATION EXTERNAL POWER AC ELECTRICAL LOAD DISTRIBUTION 2460 DC ELECTRICAL LOAD DISTRIBUTION

26 FIRE PROTECTION 2611 DETECTION—ENGINE/APU 2612 DETECTION—AIRFRAME (Cargo Compartments) 2620 EXTINGUISHING 2630 EXPLOSION PREVENTION (B747 STP)

27 FLIGHT CONTROLS 2710 AILERON AND TAB (Includes Control Wheel) 2720 RUDDER AND TAB 2730 ELEVATOR AND TAB (Stall Warning, DCB Pitch Trim) 2740 HORIZONTAL STABILIZER 2750 FLAPS (T/E) 2760 SPOILERS 2770 GUST LOCK AND DAMPERS 2780 LIFT AUGMENTING (L/E Flaps, Slats, Slots)

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28 FUEL 2810 2820 2830 2840

STORAGE (Tanks, Cells) DISTRIBUTION (Boost Pump) DUMP INDICATING (NOT Fuel Flow)

29 HYDRAULIC POWER 2910 MAIN (A and B on B727, EDP and ADP others) 2920 AUXILIARY (RAT, Standby) 2930 INDICATING (Qty, Press, Temp)

30 ICE AND RAIN PROTECTION 3010 3020 3030 3040

AIRFOIL AIR INTAKES (Nose Cowl) PITOT/STATIC WINDOWS/WINDSHIELDS AND DOORS 3050 ANTENNAS AND RADOMES 3070 WATER LINES 3080 DETECTION

31 INSTRUMENTS 3110 INSTRUMENT AND CONTROL PANELS 3120 INDEPENDENT INSTRUMENTS (Clocks) 3130 RECORDERS (Flight) 3140 CENTRAL COMPUTERS (Wt and Bal) 3150 CENTRAL WARNING SYSTEMS 3160 CENTRAL DISPLAY SYSTEM (EICAS)

32 LANDING GEAR 3210 3220 3230 3241 3242 3250 3260 3270

MAIN GEARS AND DOORS NOSE GEAR AND DOORS EXTENSION AND RETRACTION WHEEL AND BRAKES ANTI-SKID STEERING INDICATION TAIL SKID

33 LIGHTS 3310 FLIGHT COMPARTMENT 3320 PASSENGER COMPARTMENT 3330 CARGO AND SERVICE COMPARTMENTS Revision 2

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3340 EXTERIOR 3350 EMERGENCY LIGHTING

34 NAVIGATION 3410 FLIGHT ENVIRONMENT DATA (Air Data, Mach Warn, Speed, Climb Altitude, Air Temp) 3420 ATTITUDE AND DIRECTION (ADI, HSI, Compass, Instrument Comparator) 3430 LANDING AND TAXI AIDS (Radar ALT, ILS, Localizer) 3440 INDEPENDENT POSITION MEASURING (Weather Radar, INS/IRS) 3445 GROUND PROXIMITY 3450 DEPENDENT POSITION MEASURING (DME, VOR, ADF, Transponder) 3460 FLIGHT MANAGEMENT COMPUTING (FMS)

35 OXYGEN 3510 CREW 3520 PASSENGER 3530 PORTABLE

36 PNEUMATIC 3610 3611 3612 3622

DISTRIBUTION TEMPERATURE CONTROL FLOW AND PRESSURE CONTROL INDICATING

38 WATER/WASTE 3810 3820 3830 3840

POTABLE WASH WASTE DISPOSAL (Toilets) AIR SUPPLY (Water Pump)

49 AUXILIARY POWER UNIT 4940 4950 4960 4970 4990

STARTING AIR ENGINE CONTROLS INDICATING OIL

52 DOORS 5210 PASSENGER/CREW 5220 EMERGENCY EXIT Revision 2

5230 5240 5250 5260 5270

CARGO SERVICE (Hyd/Elect Compt) FIXED INTERIOR (Flt Compt) ENTRANCE STAIRS INDICATION

53 FUSELAGE 54 NACELLES AND PYLONS 55 STABILIZERS 56 5610 5620 5630 5640

WINDOWS FLIGHT COMPARTMENT CABIN DOOR INSPECTION AND OBSERVATION

57 WINGS 71 POWER PLANT (Cowlings) 72 ENGINE (Fan Blades) 73 ENGINE FUEL AND CONTROL 7310 DISTRIBUTION (Fuel Leaks) 7320 CONTROLLING (FCU, Variable Vanes) 7330 INDICATING (Fuel Flow/Used)

74 IGNITION 75 ENGINE AIR 7510 ANTI-ICING (NOT Nose Cowl) 7520 COOLING 7530 COMPRESSOR CONTROL (Bleed Valve) 7540 INDICATING

76 ENGINE CONTROLS 77 ENGINE INDICATION 7710 7720 7730 7740

POWER (EPR, RPM) TEMPERATURE (EGT, TGT) ANALYZERS (Vibration) INTEGRATED ENGINE INSTRUMENT (Use 3250 for EICAS) 2-5

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78 ENGINE EXHAUST 7830 THRUST REVERSE

79 ENGINE OIL 7910 STORAGE 7920 DISTRIBUTION (Oil Leaks) 7930 INDICATING (Temp, Press)

MAINTENANCE RECTIFICATION CODING DEFERRED TO CODES NLO—Next LayOver HGR—HanGaR CHECK ACK—A Check BCK—B Check CCK—C Check DCK—D Check

80 ENGINE STARTING DEFERRED DUE CODES

UPDATED MAS 25 EQUIPMENT AND FURNISHINGS 2510 FLIGHT COMPARTMENT— GENERAL 2511 FLIGHT CREW SEATS 2520 PASSENGER COMPARTMENT— GENERAL 2521 PASSENGER SEATS 2522 FLIGHT ATTENDANT SEATS 2523 PASSENGER SERVICE UNITS 2524 CLOSETS AND PARTITIONS 2526 STOWAGES (Overhead Bins) 2527 FLOOR COVERING (Carpets) 2530 BUFFET/GALLEY—GENERAL 2531 GALLEY (Hot Cups and Coffeemakers Use 3830 for Sinks) 2532 GALLEY INSERTS AND TROLLEYS (Ovens) 2533 REFRIGERATORS AND CHILLERS 2540 LAVATORIES (Use 3830 for Sinks, Toilets) 2552 CARGO COMPARTMENTS 2560 EMERGENCY (Life Vests/Rafts, Slides Use 3520 or 3530 for Oxygen) 2580 CREW AND PASSENGER CALL SYSTEM

P—Parts T—Time insufficient M—Manpower E—Equipment (Tools, Tow Cars) F—Facilities (Hangar Space) X—Other (Weather, etc.) ALTERNATE POSITION REFERENCE CODES APU—Auxiliary Power Unit LBG—Left Body (Landing) Gear LMG—Left Main (Landing) Gear LWG—Left Wing (Landing) Gear NLG—Nose Landing Gear PP1—Power Plant (Engine Number) 1 PP2—Power Plant (Engine Number) 2 PP3—Power Plant (Engine Number) 3 PP4—Power Plant (Engine Number) 4 RBG—Right Body (Landing) Gear RMG—Right Main (Landing) Gear RWG—Right Wing (Landing) Gear 0500—VOIDED LOG PAGES (Due soiled, write-through, etc.) ACF619-1 (DECAL)

LOG SHEET HANDLING IMMEDIATE FLX—Pull BOTH Copies DEFERRAL—Pull YELLOW Copy DEFERRAL FIX—Pull PINK Copy

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HOW TO USE THE MAINTENANCE MANUAL DIVISION OF SUBJECT MATTER The Dash 8 Maintenance Manual is divided into six major sections: “Introduction,” “Aircraft General,” “Airframe Systems,” “Structure,” “Propeller/Rotor,” and “Power Plant.” Depending upon the manufacturer and the particular airplane type, some maintenance manuals may be broken down into fewer or more major sections, if required. Each major section is, in turn, separated into chapters, with each chapter having its own effectivity page and table of contents. Only the applicable chapters are included in any particular airplane Maintenance Manual.

•

and Operation •

Pages 101 through 200............................ Testing and Troubleshooting

•

Pages 201 through 300 Maintenance Practices

NOTE The third block is used when all subtopics of Maintenance Practices are relatively brief. Whenever individual subtopics become so lengthy as to require a number of pages, the following page number blocks are used in the Dash 8 manual:

STANDARD NUMBERING SYSTEM The numbering system identifies and segregates subject matter by chapter (system), section (subsystem), and subject (unit). The system is a conventional dash-number breakdown, and each number is composed of three elements consisting of two digits each. When referred to as a unit, the three-element number (chapter/ section/subject) is called the “chapter/section” number. The chapter/section number is located in the lower right corner of each page with the page number and date. Each system, subsystem, and unit is allocated a block number to allow for coverage of various modification states. For example, if a modified rudder is introduced, the original and modified rudders will be covered under separate subject numbers: original rudder, 27-20-11; modified rudder, 27-20-12. Figure 2-1 illustrates the basic numbering system. A page numbering system allows further readout of information for rapid retrieval. All maintenance information (topics) and blocks of page numbers are assigned to each. The following page number blocks are used in the Dash 8 manual:

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Pages 1 through 100 Description

•

Pages 301 through 400............................... Servicing

•

Pages 401 through 500............. Removal/Installation

•

Pages 501 through 600.................... Adjustment/Test

•

Pages 601 through 700 .................. Inspection/Check

•

Pages 701 through 800 ................. Cleaning/Painting

•

Pages 801 through 900 .................. Approved Repair

Each new topic of information starts with page 1, 101, 201, 301, etc., and continues within the page numbering blocks as necessary; unused page number blocks are omitted. Illustrations and tables use the same numbering system as the page block they appear in–for example, Figure 403 is the third figure in the Removal/Installation topic. If an installation requires more than one page unit, whether it is a foldout or multiple-sheet presentation, each page unit will be assigned a sheet number.

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LIST OF EFFECTIVE PAGES

NOTES

A List of Effective Pages is located at the beginning of each chapter to allow the user to determine whether the chapter is complete and if it contains the latest issue of all pages. On each list, the date quoted against each page should correspond to the date on the relevant page.

SUMMARY The Introduction to an ATA 100-format Maintenance Manual provides an explanation of the documentation procedures used for that particular airplane manual. Although the ATA 100 major chapter/section/subject specification does not vary a great deal among aircraft manufacturers, there are usually minor differences among maintenance manuals, particularly in the area of techniques for marking changes and their dissemination, which do vary from one manufacturer to another. That is why the maintenance technician should read and understand the Introduction to the manual for any particular airplane which requires maintenance. The Introduction is the key to understanding manufacturer-peculiar techniques used in the manual to help the maintenance technician find the required information in a minimum of time. Review it occasionally; it has been put there to help you.

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CHAPTER/SECTION NUMBERING SECOND ELEMENT FIRST ELEMENT

THIRD ELEMENT

21-20-01 CHAPTER (SYSTEM) AIR CONDITIONING

SUBJECT (UNIT) BLOWER SECTION (SUBSYSTEM) DISTRIBUTION COVERAGE

EXAMPLES:

21-00-00 SYSTEM AIR CONDITIONING

SUBSYSTEM TEMPERATURE CONTROL

SUB-SUBSYSTEM COOLING TEMPERATURE CONTROL

21-60-00 21-62-00 21-62-03

UNIT MANUAL SHUTOFF SWITCH

WHEN CHAPTER (SYSTEM) ELEMENT NUMBER IS FOLLOWED BY ZEROS IN SECTION AND SUBJECT ELEMENTS, INFORMATION IS APPLICABLE TO THE ENTIRE SYSTEM. WHEN SECTION (SUBSYSTEM) ELEMENT NUMBER IS FOLLOWED BY ZEROS IN SUBJECT ELEMENT, INFORMATION IS APPLICABLE TO SUBSYSTEM AS A WHOLE. THIS DIGIT REPRESENTS A SUBSUBSYSTEM: INFORMATION IS APPLICABLE TO SUB-SUBSYSTEM AS A WHOLE. INFORMATION IS APPLICABLE TO SPECIFIC UNIT (COMPONENT) OF SUBSUBSYSTEM.

PAGE 201/202 BACK OF PAGE IS BLANK. MAY 1/77 DATE OF PAGE ISSUE. FIRST PAGE OF MAINTENANCE PRACTICES

NOTE: THIS FIGURE IS FOR ILLUSTRATIVE PURPOSES ONLY; IT DOES NOT REFLECT ACTUAL DASH 8 SYSTEM COMPONENT BREAKDOWN.

Figure 2-1. ATA 100 Numbering

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DASH 8 SCHEDULED MAINTENANCE PROGRAM The MSG (Maintenance Steering Group) started with the 747 and was based on the premise that scheduled overhaul does not improve reliability. The Dash 7 program used MSG-2. The Dash 8 Scheduled Maintenance Program was derived using MSG-3 logic analysis techniques which were developed by the ATA Maintenance Steering Group (MSG). MSG-3 is split into two sections, Systems and Structures.

SYSTEMS The critical items (Maintenance Significant Items or MSI’s) on the Dash 8 were identified and a MSG-3 analysis was performed on each item. The purpose of the MSG (Maintenance Steering Group) is to identify each Maintenance Significant Item, its function, potential functional failures, failure effect and failure cause. Each functional failure and failure cause is processed through logic flow charts to determine if the task is necessary and effective to maintain the designed level of safety and reliability or to provide economic benefit. The resultant tasks form part of the scheduled maintenance program.

For damage tolerance evaluation, the primary structure is divided into zones and based on crack growth analysis and determination of detectable crack size, threshold and repeat intervals are established for each Structurally Significant Item (SSI).

AIRWORTHINESS LIMITATIONS This section includes Systems and Structure tasks where the interval must be strictly controlled in accordance with the certification documentation for the aircraft. System’s limitations are normally at intervals larger than those in the MRB Report so that normal program escalation can be permitted. All fatigue inspection tasks are considered Airworthiness Limitations.

NOTES

STRUCTURES Prior to the development of the “damage-tolerant” theory, aircraft were classified by the “fail-safe” or “safe-life” method. Problems with the “safe-life” and “fail-safe” methods prompted the change to the “damage-tolerant” theory. The Dash 8 is one of the first aircraft to be certified to these damage tolerant rules. Three types of structural damage were considered and inspections were developed for each. Accidental damage (C Check), Environmental damage (Calendar) and Fatigue damage (Cycles).

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RELATED DOCUMENTATION

NOTES

The Dash 8 Maintenance Program is contained in PSM 1-8 (or 83) -7 and consists of three volumes. Volume 1 contains Part 1 MRB report and Airworthiness Limitations L i s t ( AW L ) . Vo l u m e s 2 a n d 3 c o n t a i n Supplementary Information. Part 1 MRB Report details scheduled maintenance tasks and their frequency and is approved by the MRB. Part 2, AWL, is approved by Transport Canada, Engineering. Part 3 consists of supplementary information including task procedures and details of zone and access panel numbering. The Equalized Maintenance Program (EMP) was designed to distribute the scheduled work load in such a manner that extended out-ofservice periods are avoided. This is contained in PSM 1-8 (or 83) -7E.

IMPLICATIONS OF REPAIRS TO DAMAGE TOLERANT STRUCTURE When the structure is changed in any way from certification standards, it is important to determine whether damage tolerance certification is affected by the proposed change. The two most likely ways this could occur are: 1. The proposed change may reduce fatigue life 2. The proposed change may render the certified inspection technique ineffective In case of a repair, if possible, a scheme should be chosen that does not affect the basis of certification. If this is not feasible, an SSI should be raised for the deviant aircraft. This must be evaluated by the Stress Department and revised inspection times and/or procedures specified for that particular SSI. This evaluation and the proposed inspection must be reviewed and approved by the certification authority.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TIME LIMITS/ MAINTENANCE CHECKS This maintenance procedure is for the inspection of the aircraft after a bird strike. Examine the exterior surfaces of the airplane structure in the general vicinity of the bird impact. •

•

In the event the airplane has sustained a bird strike (reported by the flight crew, or if damage is observed during routine checks) general inspection is made of the airplane to determine the areas of the bird strike, then a thorough check is made to those specific areas to determine damage that has occurred. If the initial inspection indicates structural damage, then the interior structure and any hydraulic, pneumatic, or other systems should be inspected for further damage.

NOTE: The following statements are made for the airplane structure in general and inspection can be limited to the general area of impact. Damage to the aircraft from a bird strike usually occurs in the areas that follow: • Antennas • Engines • Fuselage • Horizontal stabilizer • Landing gear • Lights • Nacelle • Nose radome • Vertical stabilizer • Windshields and side windows • Wings. It is possible for the aircraft to fly into a large flock of birds and have many bird strikes at the same time. Examine each bird strike for damage. After the flight crew gives a report of a bird strike,

TR # 4.1—January 2010

complete the inspection of the aircraft before the subsequent flight. Lightning Strike Incident Reporting To achieve a better understanding of the effects of a lightning strike on aircraft structures and systems, it is important that any lightning strike incidents are reported to Bombardier Technical Support and Engineering Departments. Please include as much information as possible regarding the lightning strike incident including:

1. Entry point of the strike. 2. Exit point of the strike. 3. Any swept stroke or intermediate attachment point. 4. Description of any physical damage or 5. electrical upset. 6. NOTE: Please provide precise data of the physical damage or electrical upset. 7. Pilot’s Report. 8. Strip report of damaged equipment. 9. Photographs of damaged areas. 10. Aircraft Information • Airline Name. • Aircraft Registration. • Aircraft Fleet No. • Aircraft Serial No. • Report Prepared by: • Name. • Title. • Phone Number.

11. Geographical Location of Strike • Aircraft Altitude: • Closest Airport: • WX: • Date: • Time: • Flight No.:

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Lightning Strike Incident Reporting (Cont) • From: • To: • Parked: • Taxi: • Take−off: • Climb: • Cruise: • Decent: • Approach: • Final: • Land:

Regarding static grounding of fuel panels on wing This will give metal to metal interfacing provides lightening protection The aircraft has many properties to limit the damage caused by a lightning strike. The primary structural components of the fuselage are fabricated from aluminum which provides a good protection for both the direct and indirect effects of a lightning discharge. The basic metal structure also supplies a ‘protective cover’ around the passengers, crew, fuel, and avionics systems B. The aircraft lightning protection systems prevent the items that follow:–Damage of the external skin, fuel tanks and other important flight safety areas–Prevent damage to low conductivity areas such as fairings and control surfaces–Prevent arcing and limit the movement of electrical charges on the external surface of the aircraft–Protect the electrical/electronic hardware on the aircraft from damage. C. Lightning protection devices are used to stop the high voltages and currents of a lightning strike. A low resistance path along the

TR4.1-2

aircraft external surface is provided for flow of electrical current between the structural components and the metal airframe. D. Lightning protection on the aircraft is provided by: 1. Metal−to−metal interfacing of fuel system components mounted on wing surfaces. 2. Electric/electronic system protection in the form of electric conduits, aluminum covers and bonding jumpers. 3. Grounding of low conductivity composite structures and metal sections to the basic metal structure 4. Bonding of control surfaces and flaps with bonding jumpers at all hinges 5. Installation of lightning diverter strips on the nose radome and application of conductive paints on nonconductive surfaces 6. Application of aluminum paste (Alcoa No. 1593 or No. 726 conforming to ASTMD962, Type 4, Class A to the primer and enamel paint on both upper and lower wing skin surfaces between stations YW171.20 and YW261.00.

TR # 4.1—January 2010

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TIME LIMITS/ MAINTENANCE CHECKS (CONT)

NOTES

DETAILS OF MAINTENANCE PROGRAM Line “L” Check: The line “L” check is to be repeated at every 50 flight hours. “A” Check: The “A” check is to be repeated at every 500 flight hours. “C” Check The “C” check is to be repeated at every 5,000

flight hours.

DIMENSIONS AND AREAS GENERAL Chapter 5 of the Maintenance Manual contains lists of airplane dimensions and areas. Included in the chapter are airplane staions, zones, and access panels. Figure 4-2 shows the general dimensions of the Series 100A and Series 300A airplanes. NOTE This information is also available at the beginning of the PSM 1-8-3 Structural Repair Manual.

TR # 4.1—January 2010

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TOWING AND TAXIING (CONT) Moving Aircraft with Flat Tires An aircraft with a flat tire or a combination of flat tires on the runway, taxiway or apron is allowed to be moved by taxiing or with the use of a tow bar or using suitable rope to clear the runway or a taxiway. The maximum combination of flat tires is defined in AMM. Make sure that the rims of the wheel with the flat tires are not damaged and the aircraft is still supported by its own main and nose landing gears.

LEVELING AND WEIGHTING (CONT) The basic weight of the aircraft is that weight which includes all fixed operating equipment both standard and optional, trapped and unusable fuel and full oil.

1. Make sure the nose landing gear ground lock is engaged. 2. Make sure the main landing gear lockpins are installed.

Obey the following precautions when towing an aircraft with two flat tires on one or more axles: • Use rope on both main landing gear for towing. • Limit towing speed to 2 MPH (3 KPH). • Towing distance should be kept to a minimum to clear the runway. • Avoid sharp turns, abrupt starts and stops.

TR # 4.2—January 2010

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PARKING AND MOORING (CONT) Note: To reduce any pitching tendency, it is recommended that the aircraft be fully fuelled and the center to gravity moved as far forward as possible. A. When parking on dry asphalt or concrete, the aircraft should be moored if wind speeds of 55 kn (102 kph) ( 63 mph) or above are expected. B. When parking on wet asphalt or concrete, the aircraft should be moored if wind speeds of 40 kn (74 kph) (46 mph) or above are expected. C. When parking on ice or snow covered surfaces, the aircraft should be moored at all times. If possible, the ice or snow should be removed from the surface underneath the tires to reduce skidding.

Note 1. Park aircraft with nose wheel centered. 2. Park aircraft heading into wind with flaps up, apply the parking brake, engage gust lock, engage nose gear ground lock, install main gear ground lock pins, install protective covers and propeller restraints, as required, close all doors and statically ground the aircraft. Before leaving the aircraft ensure that battery switches are set to OFF. Place suitable wheel chocks in front and behind the main and nose wheels. The front and rear chocks at each wheel position should be tied together for greater security. 3. For fast turn around or short term parking, apply parking brake. Place suitable wheel

TR # 4.3— March 2010

chocks at main wheels as required by local conditions

Parking Brake 1. Engage the parking brake by pulling the striped parking brake lever to the PARK position. 2. Check the brake system accumulator pressure gage located in the right nacelle to ensure sufficient pressure (2500 psi minimum). Pump up accumulator pressure with hand pump as required.

Gust Lock 1. Lock the control surfaces by setting the CONT LOCK lever to the LOCK position. 2. Set the ailerons to the neutral position and ensure that the aileron locks engage. Push either control column fully forward to engage the elevator locks. 3.

Landing Gear Ground Locks

1. Engage nose gear ground lock by pressing button on door and pulling the striped nose gear ground lock door out and turning 90 degrees clockwise. 2. Install main gear ground lock pins.

Covers 1. Engine air intake 2. Engine exhaust duct 3. Pitot head

Propeller Restraints Two propeller restraints are provided to prevent windmilling of the propellers during parking. Each restraint is secured to a fitting on the associated engine nacelle.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Static Ground Points

(c) up to 90 days

Attach one end of the grounding cable to a ground lug or other applicable grounding location. Ground points must be clean for good continuity. Attach the other end of the grounding cable to the GROUND (EARTH) HERE point on one of the two main landing gear cross beams.

(d) up to 180 days

Storage Schedules General Aircraft scheduled for storage shall be in an airworthy condition prior to being stored. Aircraft scheduled for storage shall have calendar inspection tasks which become due, while in storage, carried over and completed, prior to aircraft being returned to revenue service.

This document is designed to provide complete check list procedures for: 1. A specific target time frame i.e. “up to 90 days”. 2. To cover return to service procedures for all environmental and operational conditions. The number of check list items increase as the time in storage increases. It is only necessary to select the specific RECORD OF STORAGE i.e. “up to 180 days” the target storage time period to

(e) more than 180 days Aircraft stored for up to ninety days are considered to be in short term storage. Aircraft stored for greater than ninety days are considered to be in long term storage. •

Prior to parking the aircraft for this storage period, carry out the following functional checks:

•

Run both engines with propellers unfeathered for five minutes or until the oil temperature is in the green arc (approximately 70 degrees).

•

Select both engine bleeds to ON to operate the Environmental Control System (ECS). Cycle temperature controls to ensure positive movement of the pack and trim valves.

•

Energize AC and DC systems and ensure thesesystems are functioning properly.

•

Function deice boot system − ensure proper operation.

•

Function propeller heating system − ensure proper operation.

NOTE: One cycle at 785 RPM is acceptable. However,one or more cycles will require propeller RPM of 900 to ensure proper cooling. •

Function rudder through range of full motion − ensure proper operation (perform 2 full cycles).

Storage Time Periods

•

Schedules for initial tasks, weekly and monthly recurring tasks for the following storage time periods are provided in subsequent sections.

Function ailerons through full range of motion − ensure proper operation (perform 2 full cycles).

•

Function elevators through full range of motion ensure proper operation (perform 2 full cycles).

•

Function flaps through full range of motion − ensure proper operation (perform 2 full

(a) up to seven days (b) up to 28 days

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

•

cycles). Shut down all systems.

•

Shut down both engines.

•

Conduct a thorough check of the aircraft to ensure noleaks are present.

•

Secure aircraft once all checks are complete.

TR # 4.3— March 2010

NOTES

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FUEL SYSTEM (CONT)

NOTES

WARNING: MAKE SURE THAT THE FUEL TENDER AND THE FUEL NOZZLE ARE ELECTRICALLY BONDED TO THE AIRCRAFT BEFORE YOU DEFUEL OR REFUEL THE AIRCRAFT. AN ELECTROSTATIC SPARK DURING REFUELING/DEFUELING CAN CAUSE AN EXPLOSION OR A FIRE. NOTE: If ac power is not available, defueling will be by suction from the fuel tender only (10 in Hg maximum suction). Auxiliary pumps will not operate and pump advisory lights will not come on. WARNING: IF YOU USE THE AUXILIARY FUEL PUMP TO DEFUEL THE FUEL TANK, YOU MUST TURN THE AUXILIARY FUEL PUMP OFF AT A REMAINING FUEL QUANTITY OF 50 LBS. A FUEL QUANTITY OF 50 LBS IS SUFFICIENT TO PREVENT PUMP DRY− RUN OPERATION. DRAIN THE REMAINING FUEL WITHOUT THE USE OF THE AUXILIARY FUEL PUMP. IF YOU DO NOT DO THIS, YOU CAN CAUSE AN UNSAFE CONDITION AND POSSIBLY CAUSE AN IGNITION SOURCE INSIDE THE FUEL TANK IF AN AUXILIARY FUEL PUMP ELECTRICAL OR MECHANICAL FAILURE OCCURS. NOTE: Minimum refueling pressure is 20 psi and maximum is 50 psi.

TR # 4.4— March 2010

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EQUIPMENT & FURNISHING Each passenger service unit (PSU) incorporates a passenger service panel at each passenger seat location. Above the PSUs, overhead stowage bins are provided at each side of the aircraft for the full length of the compartment. The rear face of the wardrobe compartment wall, adjacent to the passenger door, incorporates the first flight attendant’s seat. The front face of the rear baggage access door incorporates a second flight attendant’s folding seat.

DESCRIPTION Sidewall Coverings The sidewall coverings consist of double reveal panels, sidewall panels and dado panels. The panels are of a honeycomb core, sandwiched between multi−ply layers of pre−impregnated fabric and covered with vinyl. All panels are secured to the structure by trim rails at the top and bottom.

Overhead Bins There are eight pairs of full size, one single and two half overhead bins, located on the outboard ceiling with 9G’s hanger struts. The bins are installed in pairs except the single bin which is located at the forward right side and two half bins at the rear right and left side of cabin. Each bin has a door which opens upwards. The self−latching door pivots at the front and rear and door travel is governed by gas springs, attached to the bin sidewall structure. Each pair of bins accommodates

TR # 4.5— March 2010

fluorescent ceiling lights inboard, four air duct plenums with outlet louvers and three overhead light ballasts, attached at the rear by screw and washers. The load limit for each full size bin, single bin and half bin is 60, 40 and 22 pounds respectively. The bins are constructed from a honeycomb core, sandwiched between multi−ply layers of pre−impregnated fabric and covered with vinyl. The support brackets and bin hangers are attached to the aircraft structure through anti−vibration mounts. These absorb movement from the aircraft structure and prevent vibration of the overhead stowage bins.

Passenger Seats The passenger compartment seats are mounted facing forward, except the first pair on the right side which is facing rearward. The seats are mounted on the inboard and outboard seat rails. The passenger seats are installed in pairs. The passenger seat backrests fold forward except the seats fore and aft of the Type III emergency exits. When a forward force of approximately 30 pounds is applied to the backrest of forward facing seats, a break over mechanism allow the backrest to fold forward on the bottom cushion to facilitate cleaning the aircraft interior. equipped with short outboard arm rests to aid passengers to exit the compartment, in an emergency. Literature pouches and fold away meal trays are incorporated in the backrest of seats except

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EQUIPMENT & FURNISHING (CONT) front row seats and those immediately forward of the rear divider bulkhead. For passenger seats where a meal tray is not provided provision is made for a portable tray to be plugged into the front section of the arm rests. When not in use, the portable tray are stowed in a compartment located forward of the front row aft facing seats. Passenger seats are constructed from aluminum structure, plastic trim and foam cushion. Foam bottom flotation cushions, covered with fire resistant material, are provided and the seats are fitted with shin guards, ash trays, head rest cover and lap belts.

Flight Attendant’s Annunciator Panels Two flight attendant’s annunciator panels are installed in the passenger compartment. One is located on forward ceiling panel near the passenger door and the second on the rear ceiling panel above the second flight attendant seat

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLUID SYSTEM IDENTIFICATION TAPES

FUEL LUBRICATION

YELLOW

RED

AIR CONDITIONING

BROWN GREY

FIRE PROTECTION

BROWN

INSTRUMENT AIR

HYDRAULIC

BLUE YELLOW

ORANGE GREY

DE−ICING

0.25 in. (6.35 mm)

GREY

COMPRESSED GAS

ORANGE

WATER INJECTION

RED RED GREY

BREATHING OXYGEN

GREEN

PNEUMATIC

ORANGE BLUE

FLUID SYSTEM IDENTIFICATION TAPES

TR # 4.7— March 2010

dam03_2006040_001.dg, kg, mar24/2009

1.00 in. (25.40 mm)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ELECTROSTATIC DISCHARGE SENSITIVE DEVICES − MAINTENANCE PRACTICES GENERAL Many electronic line replaceable units (LRU’s) contain micro−circuits and other sensitive devices which can be damaged internally by electrostatic discharges. These units are identified as being Electrostatic Discharge Sensitive (ESDS). Decals are installed on ESDS LRU’s to indicate that special handling is required. Personnel who remove, install, and transport ESDS LRU’s should have an understanding of static electricity including its generation and the protection required from static discharges. Damage to internal components of an ESDS LRU can be a catastrophic failure caused by a single static discharge. System characteristic changes and/or performance degradation can be caused by multiple static discharges over a long period of time. Another mode of failure does not require contact to the LRU by a static charged person or object. Simple exposure of the LRU to the electrical field surrounding a charged object can damage or degrade the LRU.

Removal/Installation of ESDS Printed Circuit Boards Removal of Printed Circuit Boards with STATIC SENSITIVE Decals

2. Prepare for printed circuit board removal as follows:

(a) Lay out the 10 mil conductive work surface adjacent to the trouble location and connect alligator clip to a suitable grounding point.

WARNING: ONLY WRIST STRAPS WITH A GROUNDING LEAD RESISTANCE OF ONE MEGOHM OR GREATER SHOULD BE UTILIZED. INADVERTENT CONTACT BETWEEN A LOW RESISTANCE WRIST STRAP AND A HIGH VOLTAGE CONSTITUTE A SHOCK HAZARD TO PERSONNEL.

CAUTION: WRIST STRAP MUST BE WORN WITH GROUNDING LEAD IN CONTACT WITH SKIN IN ORDER TO PROVIDE REQUIRED GROUNDING. FAILURE TO PROVIDE REQUIRED GROUNDING CAN RESULT IN DAMAGE TO ESDS PRINTED CIRCUIT BOARDS.

(b) Attach the wrist strap ensuring a snug fit for good contact between skin and strap. Connect alligator clip to a suitable grounding point. 3. Gain access to the equipment and determine the exact location of the printed circuit board to be removed.

1. Remove system electrical power.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

4. Use the top and bottom (or left and right) extractors on the printed circuit board to remove it.

CAUTION: DO NOT USE STAPLES OR ADHESIVE TAPES TO CLOSE CONDUCTIVE BAGS. DAMAGE TO THE CONTAINER WILL EXPOSE THE CONTENTS TO ELECTROSTATIC DISCHARGES.

5. Immediately insert the static sensitive printed circuit board into a conductive bag identified with an ESDS label. Use an ESDS label or 100% cotton twine to close the conductive bag. 6. Static sensitive printed circuit boards within a conductive bag should be placed into a rigid container in order to maintain the integrity of the conductive bag during transportation.

REMOVAL/INSTALLATION OF ESDS METAL ENCASED UNITS (LRUS) GENERAL

C. Removal of Metal Encased Units with ESDS Labels 1. Remove system electrical power 2. Loosen and remove the ESDS labelled unit from the equipment rack, airframe, or panel as directed by the Removal/Installation procedure. 3. Being sure not to touch the pins in the electrical connector, check the unit to see if a static sensitive caution decal is installed near the electrical connector(s). The static sensitive decal will be labelled similar to the following:

CAUTION: ELECTROSTATIC SENSITIVE DEVICE. CONDUCTIVE CONNECTOR DUST COVER REQUIRED.

4. The presence of such a label means the unit can be damaged by an electrostatic discharge through the connector pins. Install conductive duct cover on connectors that are labelled static sensitive and standard dust covers on connectors that are not labeled. 5. Conductive dust caps and connector covers from the unit being installed may be used on the unit being removed.

Metal encased units can be rack mounted, bolted to airplane structure, or control panels installed in instrument panels.

ELECTRICAL

Equipment

Safety Procedures

Conductive electrical dust caps and connector covers

The procedure that follows will help to prevent injury to the person and damage to equipment:

NOTE: Conductive dust caps and connector covers are black or grey in color.

1. Refer to the AMM Chapter 12, for the correct procedure before you energize the DC or AC electrical system. 2. If a circuit breaker must stay open until a maintenance procedure is completed, install

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

a safety clip below the circuit breaker button. Tag the circuit breaker to tell the personnel not to close the circuit breaker. 3. If a switch must stay open or closed until a maintenance procedure is completed, tag the switch to tell the personnel to not move the switch. 4. When the maintenance work is completed, make sure that all the connectors are correctly reinstalled. Make sure that all switches and controls are in a position to prevent accidental operation of components. Remove the tags from the switches, and the clips and tags from the circuit breakers. 5. Energize the electrical system. Do the operational and functional checks. 6. Make sure that all the switches and controls are back to their correct position after the end of the operational or functional checks.

Circuit Breaker Reset 1. When a circuit breaker opens, do not try to reset until you repair the fault. 2. Close the circuit breaker only after you repair the fault.

TR # 4.8— March 2010

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Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CHAPTER 21 AIR CONDITIONING (SERIES 100/100A) CONTENTS Page INTRODUCTION................................................................................................................ 21-1 GENERAL ........................................................................................................................... 21-3 BLEED-AIR SYSTEM ........................................................................................................ 21-5 General.......................................................................................................................... 21-5 Components Description and Operation....................................................................... 21-7 Control and Indication................................................................................................ 21-17 Operation .................................................................................................................... 21-21 Electrical Failure ........................................................................................................ 21-21 Overheat ..................................................................................................................... 21-21 AIR-CONDITIONING SYSTEM ..................................................................................... 21-23 General ....................................................................................................................... 21-23 Cooling System .......................................................................................................... 21-23 Components Description and Operation .................................................................... 21-27 Operation .................................................................................................................... 21-39 TEMPERATURE CONTROL SYSTEM .......................................................................... 21-41 General ....................................................................................................................... 21-41 Operation .................................................................................................................... 21-43 CONDITIONED AIR DISTRIBUTION (PRE-1990 INTERIOR)–SERIES 100 ............. 21-47 General ....................................................................................................................... 21-47 Cabin Air Distribution................................................................................................ 21-47

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Flight Compartment Air Distribution......................................................................... 21-49 Avionics Compartment Cooling ................................................................................. 21-53 Air Recirculation System ........................................................................................... 21-53 CONDITIONED AIR DISTRIBUTION (1990 INTERIOR)—SERIES 100A................. 21-55 General ....................................................................................................................... 21-55 Cabin Air Distribution................................................................................................ 21-57 Flight Compartment Air Distribution......................................................................... 21-59 Gasper Air System...................................................................................................... 21-59 Air Recirculation System ........................................................................................... 21-61 Avionics Compartment Cooling (1990 Interior) ........................................................ 21-63 Operation .................................................................................................................... 21-63 PRESSURIZATION SYSTEM—SERIES 100 ................................................................. 21-65 General ....................................................................................................................... 21-65 Components Description and Operation .................................................................... 21-67 Operation .................................................................................................................... 21-73

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ILLUSTRATIONS Figure

Title

Page

21-1

Air-Conditioning System .................................................................................... 21-2

21-2

Bleed-Air System Schematic .............................................................................. 21-4

21-3

High-Pressure Bleed Shutoff Valve .................................................................... 21-6

21-4

High-Pressure Switch.......................................................................................... 21-8

21-5

Low-Pressure Check Valve ................................................................................. 21-8

21-6

Pressure Regulator and Shutoff Valve .............................................................. 21-10

21-7

Pressure Regulator and Shutoff Valve Schematic (Valve Closed) .................... 21-12

21-8

Pressure Regulator and Shutoff Valve Schematic (Normal Regulated Flow) ................................................................................. 21-14

21-9

Bleed-Air Overtemperature Switch .................................................................. 21-16

21-10

Bleed-Air Control and Indication ..................................................................... 21-16

21-11

Bleed-Air System.............................................................................................. 21-18

21-12

Bleed-Air Electrical Schematic ........................................................................ 21-20

21-13

Air-Conditioning System Controls ................................................................... 21-22

21-14

Air-Conditioning Pack Schematic .................................................................... 21-24

21-15

Temperature Trim Valves.................................................................................. 21-26

21-16

Pack Temperature Control Valves..................................................................... 21-28

21-17

Heat Exchanger................................................................................................. 21-30

21-18

Condenser/Mixer .............................................................................................. 21-32

21-19

Air Cycle Machine............................................................................................ 21-34

21-20

Compressor Discharge Overtemperature Switch.............................................. 21-36

21-21

Temperature Controllers ................................................................................... 21-36

21-22

Temperature Sensors and Switches................................................................... 21-38

21-23

AIR CONDITIONING Control Panel .............................................................. 21-38

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21-24

Temperature Control System ............................................................................ 21-40

21-25

Cabin Temperature Bias Control ...................................................................... 21-42

21-26

Cabin Air Distribution ...................................................................................... 21-46

21-27

Diverter System ................................................................................................ 21-46

21-28

Flight Compartment Air Distribution ............................................................... 21-48

21-29

Diverter Valve Installation (Mod 8/0563)......................................................... 21-48

21-30

Flight Compartment and Fan System ............................................................... 21-50

21-31

Flight Compartment Fan Electrical Schematic ................................................. 21-50

21-32

Recirculation Fan.............................................................................................. 21-52

21-33

Air Recirculation System.................................................................................. 21-52

21-34

Conditioned Air Distribution (1990 Interior) ................................................... 21-54

21-35

Cabin Air Distribution Schematic (1990 Interior) ............................................ 21-56

21-36

Flight Compartment Air Distribution (Mod 8/1538) ........................................ 21-58

21-37

Gasper System Schematic (1990 Interior) ........................................................ 21-58

21-38

Air Recirculation System Schematic (1990 Interior) ....................................... 21-60

21-39

Avionics Compartment Cooling Fan ................................................................ 21-62

21-40

Pressurized Areas.............................................................................................. 21-64

21-41

Pressurization Controls and Indicators ............................................................. 21-64

21-42

Normal Outflow Valve—Series 100 ................................................................. 21-66

21-43

Pressurization Control Schematic—Series 300 ................................................ 21-68

21-44

Selector Panel. .................................................................................................. 21-70

21-45

Indicator Panel .................................................................................................. 21-72

21-46

Pressurization System Schematic ..................................................................... 21-74

21-47

Pressurization Envelope.................................................................................... 21-76

21-48

Forward Dump Manual Selector....................................................................... 21-76

21-49

Pressurization Control—Electrical Schematic.................................................. 21-78

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TABLE Table 21-1

Revision 2

Title

Page

Pressurization Control Settings......................................................................... 21-72

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CHAPTER 21 AIR CONDITIONING (SERIES 100/100A)

INTRODUCTION This chapter, though titled “Air Conditioning,’’ deals with the environmental systems of the Dash 8, including bleed air, air-conditioning, and pressurization. Information is included from Chapter 21, “Air Conditioning’’ and Chapter 36, “Pneumatics,’’ of the Maintenance Manual. The material in this chapter is oriented toward the line mechanic. All values expressed throughout this chapter, such as for pressure, temperature, flow rates, and time, are used only for their illustrative meanings. Actual values may differ and must be obtained from the pertinent sections of the Maintenance Manual.

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FRONT PRESSURE BULKHEAD FLIGHT COMPARTMENT DISTRIBUTION

BACKUP PRESSURIZATION CONTROL FLIGHT COMPARTMENT BULKHEAD AIR RECIRCULATION SYSTEM BLEED-AIR SYSTEM

AIR CONDITIONING OFF

RECIRC 60 40 CABIN

20

1 BLEED 2

80

DUCT 100 0 °C MAN

COOL

WARM

F/A CABIN

AUTO

TEMP CONTROL

MIN MAX BLEED

COOL

WARM

CABIN DISTRIBUTION

REAR PRESSURE DOME

NORMAL PRESSURIZATION CONTROL

REFRIGERATION PACK

FLT COMP

AIR CONDITIONING PANEL

Figure 21-1. Air-Conditioning System

21-2

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GENERAL

NOTES

For descriptive purposes, this chapter is divided into three sections: bleed air, air conditioning, and pressurization. The first section describes the bleed-air source and its control and distribution to the various services. The second section describes the inflow control of air to the occupied areas of the airplane and the conditioning of this inflow to achieve and maintain the desired area temperature. The third section describes the outflow control of air from the occupied areas to achieve and maintain the desired cabin altitudes and rates of change conducive to maximum personal comfort. The air-conditioning system provides an inflow of temperature-controlled air to the cabin and flight compartment for heating, cooling, and ventilation. Air required for system operation is obtained from each engine through highand low-pressure bleed ports. On airplanes incorporating SOO 8062, bleed air is also available from the auxiliary power unit (APU) when the airplane is on the ground. The air-conditioning system is shown in Figure 21-1.

Revision 2

21-3

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

1

2

HP NO. 2 ENGINE 9 LP 7

TO DEICING SYSTEM

8

10

4

11

3

FOR DETAILS REFER TO PRESSURE REGULATOR AND SHUTOFF VALVE SCHEMATIC (FIGURE 21-7)

5

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

HIGH-PRESSURE SWITCH HIGH-PRESSURE SHUTOFF VALVE SERVO PRESSURE PRESSURE REGULATOR AND SHUTOFF VALVE BLEED-AIR OVERTEMPERATURE SWITCH WING ISOLATION CHECK VALVE LOW-PRESSURE BLEED CHECK VALVE CHOKING VENTURI PRESSURE REGULATOR ELECTROPNEUMATIC SERVO TORQUE MOTOR SERVO AIR FILTER

6

TO HEAT EXCHANGER

6

FROM NO. 1 ENGINE BLEED-AIR SYSTEM (SIMILAR TO NO. 2 SYSTEM)

Figure 21-2. Bleed-Air System Schematic

21-4

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

BLEED-AIR SYSTEM

NOTES

GENERAL Within each nacelle, the bleed-air system consists of a high-pressure shutoff valve, a highpressure switch, a pressure regulator and shutoff valve, a flow control servo, a bleed-air overtemperature switch, a low-pressure bleed check valve, and two choking venturis. Both engine outputs join downstream of two wing isolation check valves. A single duct routes the bleed air back to the air-conditioning pack (or packs) in the rear fuselage, aft of the rear pressure dome. Each engine bleed is electrically controlled through a BLEED switch and a high-pressure bleed control relay. A single BLEED flow control knob adjusts both pressure regulator settings and thus controls the flow rate. Power for electrical control is taken from the 28VDC left main and secondary buses through BLEED SYS CONT 1 and BLEED SYS FLOW CONT circuit breakers for the left (No. 1) bleed-air system, and from the 28VDC right main bus through a BLEED SYS CONT 2 circuit breaker for the right (No. 2) bleed-air system. Bleed-air overheat is sensed by an overtemperature switch which initiates closing of the high-pressure shutoff valve and pressure regulator and shutoff valve and illuminates a BLEED AIR caution light. A tapping upstream of the pressure regulator and shutoff valve provides an air supply for deicing purposes. The bleed-air system is shown schematically in Figure 21-2.

Revision 2

21-5

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SOLENOID SHOWN ENERGIZED A B C

FROM PRESSURE SWITCH

SOLENOID SHOWN DEENERGIZED FROM PRESSURE SWITCH

CONNECTOR

A B C

CONNECTOR VALVE CLOSED

VALVE OPEN

FILTER

FILTER

DIRECTION OF FLOW

DIRECTION OF FLOW

TWO-INCH DISC

Figure 21-3. High-Pressure Bleed Shutoff Valve

21-6

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION AND OPERATION

NOTES

High-Pressure Bleed Shutoff Valve The high-pressure bleed shutoff valve (Figure 21-3) is mounted in the hot air duct leading from the P 3 bleed port on the engine. It is normally closed, solenoid-controlled, and pneumatically operated and is used to control flow of high-pressure bleed air from the engine. The valve consists of a solenoid and a pneumatic actuator that is mechanically linked to a butterfly valve. When the solenoid is energized, upstream servo pressure operates the actuator to open the butterfly valve. To ensure correct installation, a flow direction arrow is provided on the valve body. The function of the valve is to ensure adequate airflow to the airconditioning system for passenger comfort at lower engine power settings. The valve will fail-safe closed if electrical failure occurs.

Revision 2

21-7

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 21-4. High-Pressure Switch HOUSING ASSEMBLY

SEAT

PISTON

OPEN POSITION

CLOSED POSITION

Figure 21-5. Low-Pressure Check Valve

21-8

Revision 4

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

High-Pressure Switch

NOTES

The high-pressure switch (Figure 21-4), located on the wing front spar outboard of the nacelle, is connected to the outlet duct from the engine bleed port. The pneumatic pressure-sensing switch contains diaphragm-operated contacts that are connected into the circuit to the high-pressure bleed valve when the BLEED switch in the flight compartment is selected on. Increasing pressure of 55 psi moves the diaphragm to open the switch contacts; the contacts close when the sensed pressure drops to 50 psi.

Choking Venturis The two choking venturis (Figure 21-2) are mounted in the outlet ducts of the high- and low-pressure bleed ports. The venturi in the high-pressure line restricts bleed to 10% of engine airflow; the low-pressure venturi restricts bleed to 3% of engine airflow. The venturis ensure adequate air in the compressor for high-demand situations such as single-engine operation.

Low-Pressure Bleed Check Valve The check valve (Figure 21-5) is a springloaded-closed, poppet-type valve mounted in the outlet duct from the engine low-pressure bleed port. The purpose of the valve is to prevent high-pressure bleed air flowing into the low-pressure bleed port when the system is operating on high pressure. The valve allows low-pressure flow into the system when the high-pressure shutoff valve is closed.

Revision 2

21-9

Canada Ltd.

ER LT

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FI

Figure 21-6. Pressure Regulator and Shutoff Valve

21-10

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Pressure Regulator and Shutoff Valve

NOTES

General The pressure regulator and shutoff valve (Figure 21-6), installed in the bleed-air duct in the nacelle, has two functions: to shutoff air to the air-conditioning system and to regulate the flow of bleed air into the system. It consists of a solenoid, a pneumatic actuator mechanically connected to a butterfly valve, and a pressure-sensing control to maintain a schedule of bleed airflow.

Revision 2

21-11

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

BLEED FLOW CONTROL ELECTROPNEUMATIC SERVO BLEED MAX MIN

TORQUE MOTOR FILTER

28 VDC L&R SEC MOD 8/1820

18 PSI FROM DEICE SYSTEM

PRESSURE SENSING CONTROL

SERVOACTUATOR

STABILIZER UNIT

ANEROID RESTRICTOR 28 VDC SOLENOID (SHOWN ENERGIZED)

AMB SUMMING BAR

TEST PORT

SENSING ACTUATOR

PNEUMATIC ACTUATOR

CLOSE FILTER

LEGEND

AMB

BLEED AIR REGULATED AIR REGULATED AIR (18 PSI) AMBIENT BUTTERFLY VALVE CLOSED

Figure 21-7. Pressure Regulator and Shutoff Valve Schematic (Valve Closed)

21-12

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Operation

NOTES

The pressure regulator and shutoff valve (Figure 21-7) is pneumatically operated and electrically controlled. It consists of a solenoid, a pneumatic actuator that is mechanically linked to a butterfly valve, and a pressure-sensing control to maintain bleed-air flow schedule. It is powered from the 28-VDC left and right main buses. Pressure for valve operation is from upstream of the valve and is routed to both sides of the pneumatic actuator piston. The close side is spring-assisted; the open side has a larger area to overcome the combined force of the spring and air pressure. The amount of pressure to the open side is controlled by the pressure-sensing control, which consists of four pneumatic inputs to a summing bar that controls pressure venting to establish open side pressure. These inputs are: • An actuator sensing downstream pressure • A stabilizer unit controlled by a restricted line • An aneroid that senses pressure altitude pressure changes • A servoactuator that receives the output of the flow control servo The electropneumatic servo torque motor (Figure 21-7), located adjacent to the pressure regulator and shutoff valve in the nacelle, meters deicing system pressure to the pressure-sensing control. A DC electrical signal from the bleed flow control positions a torque motor to vent servo pressure overboard.

Revision 2

21-13

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

BLEED FLOW CONTROL ELECTROPNEUMATIC SERVO BLEED MAX MIN

TORQUE MOTOR FILTER

28 VDC L&R SEC MOD 8/1820

18 PSI FROM DEICE SYSTEM

PRESSURE SENSING CONTROL

SERVOACTUATOR

STABILIZER UNIT

ANEROID RESTRICTOR 28 VDC SOLENOID (SHOWN DE-ENERGIZED)

AMB SUMMING BAR

TEST PORT

SENSING ACTUATOR

PNEUMATIC ACTUATOR

CLOSE FILTER

LEGEND

AMB

BLEED AIR REGULATED AIR REGULATED AIR (18 PSI) AMBIENT BUTTERFLY VALVE (OPEN)

Figure 21-8. Pressure Regulator and Shutoff Valve Schematic (Normal Regulated Flow)

21-14

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Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

The pressure regulator and shutoff valve is fail-safe open in the event of an electrical failure. Selecting the BLEED switch on the AIR CONDITIONING panel to OFF energizes the solenoid to close off the supply of air pressure to the open side of the actuator (Figure 21-7). The spring and air pressure to the close side moves the actuator, closing the butterfly. Selecting the BLEED switch to BLEED deenergizes the solenoid, allowing air pressure to open the butterfly valve (Figure 21-8).

valve operation is the same as previously described. To reduce the amount of air being bled from the engine as altitude increases, an aneroid expands to vent more air pressure. The aneroid contracts as altitude decreases, venting less air pressure.

NOTES

With the BLEED switch in BLEED, the valve regulates bleed-air flow to the air-conditioning system. This is accomplished by the pressure-sensing control venting a measured quantity of pneumatic actuator open-side air pressure overboard; the amount is dependent on three interrelated factors: • Bleed flow rate required • Altitude • Bleed-air pressure sensed downstream of the valve If a high flow rate is selected by the BLEED control knob, the torque motor is positioned to vent a minimum amount of pressure from the open side of the pneumatic actuator. The remaining pressure overcomes the spring to drive the butterfly toward the open position. If a low flow rate is desired, the torque motor is positioned to vent a small amount of servo pressure overboard, the remaining pressure acting on the servoactuator to vent more pressure from the open side of the actuator. Spring force overcoming the servo pressure drives the butterfly toward the closed position. With the flow rate set, air venting is controlled by pressure sensed downstream of the valve. Pressure increase acts on a sensing actuator to vent more air; decreasing pressure results in the sensing actuator returning to normal and less air being vented. The stabilizer unit dampens movement of the sensing actuator. Pneumatic actuator and butterfly

Revision 2

21-15

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 21-9. Bleed Air Overtemperature Switch #1 BLEED HOT

CAUTION LIGHTS

#2 BLEED HOT

ILLUMINATE TO WARN OF OVERHEAT CONDITION (OVER 287° C) IN RELATED BLEED-AIR SUPPLY DUCT DOWNSTREAM OF PRESSURE REGULATOR AND SHUTOFF VALVE. AFFECTED PRSV AND HIGH-PRESSURE BLEED VALVE CLOSE TO CUT OFF HOT AIR SUPPLY TO DUCT (BLEED SUPPLY TO DEICING SYSTEM VIA LOW-PRESSURE PORT IS UNAFFECTED). WHEN TEMPERATURE DROPS BELOW THE OVERTEMPERATURE LEVEL, BLEED-AIR OPERATION IN AFFECTED NACELLE IS RESTORED. BLEED-AIR CONTROL SWITCH (2) INITIATES BLEED-AIR FLOW TO AIR-CONDITIONING SYSTEM FROM ASSOCIATED NACELLE. AIR CONDITIONING

OFF RECIRC

1

BLEED

2

BLEED-AIR FLOW CONTROL 60 40 CABIN 80 20

SIMULTANEOUSLY CONTROLS PRESSURE FROM LEFT AND RIGHT PRSVS, WHICH DETERMINES AIR-CONDITIONING SYSTEM FLOW RATE.

DUCT 100 0 °C

MIN MAN

COOL

WARM

MAX BLEED

COOL

NOTE: MOD 8/1155 ADDS A DETENT POSITION AT “MIN.”

WARM

AUTO

F/A CABIN

TEMP CONTROL

FLT COMP

Figure 21-10. Bleed-Air Control and Indication

21-16

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Overtemperature Switch

NOTES

The overtemperature switch (Figure 21-9) is installed in the bleed-air delivery duct downstream of the pressure regulator and shutoff valve. It consists of a normally open, thermal switch with a bimetallic element and closes when air temperature in the duct exceeds 550 ±10° F (287 ±5° C).

CONTROL AND INDICATION BLEED Switches The two BLEED switches (Figure 21-10) on the AIR CONDITIONING panel control through the left and right DC buses the opening and closing of the HP shutoff valves and the pressure regulator and shutoff valves.

Bleed Flow Control The bleed flow control on the AIR CONDITIONING panel is a potentiometer which controls, through the left and right secondary buses, the electropneumatic servo in each nacelle. Adjustment of the control causes changes in bleed flow volume through the pressure regulator and shutoff valves.

Caution Lights Bleed-air overheat conditions are sensed by an overtemperature switch which initiates closing of the HP shutoff valve and pressure regulator and shutoff valve and the illumination of the BLEED HOT caution light on the caution lights panel.

Revision 2

21-17

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

1 2

HP NO. 2 ENGINE

9

LP 7

TO DEICING SYSTEM

8

10

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

HIGH-PRESSURE SWITCH HIGH-PRESSURE SHUTOFF VALVE SERVO PRESSURE PRESSURE REGULATOR AND SHUTOFF VALVE BLEED-AIR OVERTEMPERATURE SWITCH WING ISOLATION CHECK VALVE LOW-PRESSRUE REGULATOR VALVE CHOKING VENTURI PRESSURE REGULATOR FLOW CONTROL SERVO SERVO AIR FILTER

11

3

4

LEGEND 5

P3 AIR P2.5 AIR REGULATED AIR (DEICING) PRSV

6

TO HEAT EXCHANGER

6

FROM NO. 1 ENGINE BLEED-AIR SYSTEM (SIMILAR TO NO. 2 SYSTEM)

NO BLEED

Figure 21-11. Bleed-Air System (Sheet 1 of 2)

21-18

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

1

1

2

2 HP

HP

NO. 2 ENGINE LP

NO. 2 ENGINE LP

9

7

TO DEICING SYSTEM

8 10 11

9

7

TO DEICING SYSTEM

8

3

10 11

3

4

4

LEGEND P3 AIR P2.5 AIR

5

5

REGULATED AIR (DEICING) PRSV 6

6

TO HEAT EXCHANGER

TO HEAT EXCHANGER 6

6

FROM NO. 1 ENGINE BLEED-AIR SYSTEM (SIMILAR TO NO. 2 SYSTEM)

FROM NO. 1 ENGINE BLEED-AIR SYSTEM (SIMILAR TO NO. 2 SYSTEM)

LOW BLEED

HIGH BLEED

Figure 21-11. Bleed-Air System (Sheet 2 of 2)

Revision 2

21-19

21-20 HIGHPRESSURE SWITCH

HIGH-PRESSURE FLOW SHUTOFF CONTROL VALVE SERVO

PRESSURE REGULATOR AND SHUTOFF VALVE

BLEED AIR OVERTEMPERATURE SWITCH

HIGHPRESSURE SWITCH

PRESSURE REGULATOR AND HIGH-PRESSURE FLOW SHUTOFF SHUTOFF CONTROL VALVE VALVE SERVO

BLEED AIR OVERTEMPERATURE SWITCH

1 A2

A3 A1 B3

B2

B1

LOGIC CIRCUIT

NO. 1 BLEED HOT

LOGIC CIRCUIT

AIR CON PACK HOT

LOGIC CIRCUIT

NO. 2 BLEED HOT

9811-TB16

X1 A3

A3

A2

X2

A1 B3

B2

A1

H

B3 X1

B1

J

X2

K1 HP BLEED CONT RELAY 1 BLEED CONT 1 SW

X1 X2

K1 HP BLEED CONT RELAY

3 ON

1 BLEED CONT SW 1

3

OFF

ON 2

SEE CHAP 30 5

DEICE SYSTEM LP WARN SW

28V DC L MAIN

A

G

B

A2

A3

B2

A1 B3 B1

D COMPRESSOR DISCHARGE OVERTEMPERATURE SWITCH CLOSES WHEN TEM > 207°C

X1

OFF

DEICE SYSTEM SEE LP CHAP 30 WARN SW BLEED SYS 5 CONT 1 2

C

F

(SPARE) K E

B1 B2

K2 HP BLEED CONT RELAY

9811-TB16

B A

A2

A2

A3 A1

C1 X1 X2

R

3010K2

3

TIME DELAY ON OPERATE 6 SEC

X2

R1

K3 MOD 8/0496 TIME DELAY ON RELEASE

BLEED FLOW CONTROL 28V DC L SEC

ON R2

NOTE: IDENT CODE IS 2121 UNLESS OTHERWISE INDICATED.

1 OFF 2 BLEED CONT SW 5 BLEED SYS CONT 2 28V DC R MAIN

BLEED SYS CONT 1

28V DC L MAIN

Revision 2

Canada Ltd.

MOD 8/1375

MASTER CAUTION PANEL

28V DC

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

1

A B

C B A

B A C

A B C

A B C

A B

C B A

A B C

A B C

B A C

CLOSES WHEN TEMP > 287°C

OPENS WHEN PRESSURE > 55 PSIG

Figure 21-12. Bleed-Air Electrical Schematic

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

OPERATION

OVERHEAT

Selecting the BLEED switch on the AIR CONDITIONING panel to BLEED initiates two electrical functions in the bleed-air system. The solenoid of the high-pressure bleed valve is energized open through the normally closed contacts of the HP bleed control relay K1 and the high-pressure switch. The pressure regulator and shutoff valve is deenergized open. With the high-pressure bleed valve open and an engine operating, bleed air flows to the airconditioning system through the choking venturi and to the pressure regulator and shutoff valve (Figure 21-11), which regulates the air to approximately 35 psi. When bleed-air pressure reaches 55 psi, the high-pressure switch opens to interrupt the circuit to the high-pressure bleed valve, which then closes. Low-pressure bleed air then supplies the system. If high bleed-air pressure drops to 50 psi, the highpressure switch closes, opening the high-pressure bleed valve. If Mod 8/0350 is incorporated, the high-pressure valves will not open until both engines have experienced a pressure drop.

If an overheat condition (550° F [287° C]) closes the overtemperature switch, three actions are initiated as the HP bleed control relay is energized: (1) the circuit to the high-pressure bleed valve is interrupted to deenergize the valve closed, (2) the circuit to the pressure regulator and shutoff valve is completed to energize the valve closed, and (3) the BLEED HOT caution light on the caution lights panel illuminates. When the overheat condition clears, the system resets to function normally.

NOTE An overheat condition in the bleedair system from one engine does not affect the operation of the bleed-air system from the opposite engine.

NOTE Mod 8/1375 allows the HPSOVs to remain open, with the bleed switches off, so that a higher pressure air can be supplied to the deicing system in normal ground operations.

ELECTRICAL FAILURE If an electrical failure occurs with the BLEED switch in BLEED, the high-pressure bleed valve is deenergized closed to prevent thermal damage from the hot bleed air. The pressure regulator and shutoff valve must be energized closed; therefore, it remains open when electrical power is off. In this mode, low-pressure bleed air is supplied to the air-conditioning system for temperature control and to maintain cabin pressurization (Figure 21-12).

Revision 2

21-21

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RECIRCULATION FAN SWITCH ACTIVATES RECIRCULATION SYSTEM FAN WHEN SELECTED TO RECIRC.

AIR CONDITIONING OFF

CABIN DUCT TEMPERATURE INDICATOR INDICATES AIR TEMPERATURE IN CABIN DISTRIBUTION SUPPLY DUCT.

60 40 CABIN

20

0

MAN–AUTO SWITCH (2) AT MAN, TEMPERATURE CONTROL OF CABIN OR FLIGHT COMPARTMENT IS BY RELATED COOL–WARM SWITCH. AT AUTO, CONTROL IS DIRECTED TO ASSOCIATED AUTOMATIC TEMPERATURE CONTROLLER.

OFF

RECIRC

1

BLEED

MIN

WARM

NOTE: F/A POSITION AT MAX COOL ON THE CABIN RHEOSTAT MOD 8/0807

MAX BLEED

COOL

WARM

AUTO

F/A CABIN

CABIN COOL–WARM SWITCH WITH RELATED MAN–AUTO SWITCH AT MAN, DIRECTLY CONTROLS PACK TEMPERATURE CONTROL VALVES WHEN HELD AT SPRING-LOADED COOL OR WARM POSITION.

FC FAN

80

DUCT 100 °C MAN

COOL

2

FLIGHT COMARTMENT FAN SWITCH ACTIVATES FAN IN AIRCRAFT NOSE TO VENTILATE FLIGHT COMPARTMENT WHEN ON THE GROUND. FAN IS DISARMED WHEN AIRBORNE. (PRE-1990 INTERIOR)

TEMP CONTROL

FLT COMP

FLIGHT COMPARTMENT COOL–WARM SWITCH WITH RELATED MAN–AUTO SWITCH AT MAN. DIRECTLY CONTROLS TEMPERATURE TRIM VALVES WHEN HELD AT SPRING-LOADED COOL OR WARM POSITIONS. AUTOMATIC TEMPERATURE SELECTOR (2) ROTARY KNOB, WHEN POSITIONED AS DESIRED, INPUTS TEMPERATURE REQUIREMENT INTO AUTOMATIC TEMPERATURE CONTROLLER, WHICH MANIPULATES PACK TEMPERATURE VALVES OR TEMPERATURE TRIM VALVES TO MAINTAIN THE SELECTED LEVEL (MAN–AUTO SWITCH MUST BE AT AUTO.)

Figure 21-13. Air-Conditioning System Controls

21-22

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR-CONDITIONING SYSTEM GENERAL The air-conditioning system provides an inflow of temperature-controlled air to the cabin and flight compartment for heating, cooling, and ventilation. The system consists of a bleed-air system for each engine, an air-conditioning pack, and an air recirculation system. Conditioned air is supplied to the cabin and flight compartment through outlet grilles in the cabin ceiling and/or the lower sidewall vents flight compartment sidewalls. With the recirculation system operating, air is supplied to individually controlled outlets in the cabin, flight compartment, and lavatory. Bleed air from the engines is hot enough to meet all cold day requirements. On hot days, the bleed air is routed through an air cycle machine, where it is cooled to below ambient, using the pressure and energy of the air to drive the cooling turbine. When the air is cooled below its dew point, condensate is extracted by a condenser.

The cabin system ducts the air up to and along the cabin ceiling (or via the underfloor ducting depending on cabin air supply switch position), where it is discharged through lowvelocity air outlets along each side of the center aisle. The flight compartment system ducts air forward under the cabin floor to flight compartment side window demist outlets, directionally adjustable side console outlets, and fixed outlets at the rudder pedal foot wells. The air recirculation system improves cabin airflow and reduces the flow demand required of the air-conditioning system. Air is drawn from the flight compartment, through the avionics compartment, and into the underfloor duct by an electric recirculation fan in the duct, which forces the air aft to the mixer. Branch ducts just downstream of the fan direct part of the recirculation flow to a series of high-velocity air outlets (gaspers) located above each passenger position, in the lavatory, and on the pilot’s and copilot’s side panels. Air-conditioning system controls are shown in Figure 21-13.

NOTES COOLING SYSTEM General The major component of the cooling system is an air-conditioning pack located in the aft fuselage. It is supplied with bleed air from both engines and cools this air as necessary before distribution to the cabin and flight compartment. When heating is required, bleed air bypasses the refrigeration section of the pack. If cooling is required, the bypass valve is closed to divert bleed air through the refrigeration section. Temperature of the air supplied by the pack is controlled either automatically or manually from the flight compartment or the flight attendant’s panel if the cabin system is selected to AUTO.

Revision 2

21-23

21-24

AIR CONDITIONING OFF RECIRC

OFF 1

BLEED

2

FC FAN

60 40 CABIN 80 20 0

DUCT 100 °C

MIN MAN

COOL

WARM

MAX BLEED

COOL

WARM

AUTO

CABIN DUCT HOT

FLT COMPT DUCT HOT "FULL COLD" OVERRIDE SIGNAL

CABIN TEMP BIAS CONTROL

TEMP CONTROL

F/A CABIN

FLT COMP

FLIGHT COMPARTMENT TEMPERATURE CONTROLLER

CABIN TEMPERATURE CONTROLLER

FLIGHT COMPARTMENT ZONE SENSOR INPUT

"FULL COLD" OVERRIDE SIGNAL TO RAM-AIR SCOOP

DUCT GAUGE TEMP SENSOR

START BYPASS

BL AIR

DEICE SYSTEM

OPEN

BAFFLE BOX

TRIM VALVE ACTUATOR TEMPERATURE TRIM VALVES BLEED-AIR SUPPLY DUCT

L ENGINE BLEED AIR

DUCT TEMP SENSOR

APU

PACK TEMPERATURE CONTROL VALVES

ACM SUPPLY DUCT

APU BLEED-AIR VALVE

CONTROL VAVLE ACTUATOR

R ENGINE BLEED AIR

RAM AIR OVERBOARD

CABIN DUCT

LEGEND

DUCT OVERTEMP SWITCH

BLEED AIR FLT COMPT DUCT DUCT TEMP SENSOR

RECIRCULATED CABIN AIR

COMPRESSOR DISCHARGE OVERTEMPERATURE SWITCH

HEAT EXCHANGER AIR CONDENSER

COLD AIR

MIXING BOX WATER FILTER

HEAT EXCHANGERS

WATER NOZZLE

RECIRCULATED AIR CABIN CONDITIONED SUPPLY

WATER TRAP

FLIGHT COMPARTMENT CONDITIONED SUPPLY

FAN

AMBIENT OR RAM AIR

COMPRESSOR

WATER

Revision 2

AIR COND PACK HOT

EXPANSION TURBINE

ACM BYPASS DUCT

RAM-AIR DUCT

SHUTDOWN SIGNAL TO BLEED-AIR SYSTEMS

Figure 21-14. Air-Conditioning Pack Schematic

ELECTRICAL POWER

Canada Ltd.

CAUTION LIGHT

AIR CYCLE MACHINE (ACM)

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN ZONE SENSOR INPUT

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Air-Conditioning Pack The air-conditioning pack (Figure 21-14), located in a bay in the aft fuselage, consists of an air cycle machine (ACM), primary and secondary heat exchangers, an exhaust duct, and associated piping and control valves. The pack is supplied with hot bleed air ducted in along the dorsal fin. According to temperature requirements, this air is cooled or passed as hot air to the mixer box.

Conditioned air leaving the mixer/condenser is separated by a diverter baffle for entry into the cabin and flight compartment air distribution systems.

NOTES

Air to be cooled is ducted through the primary heat exchanger, reducing its temperature, before entering the compressor side of the ACM where its pressure is increased. The compressor discharge is cooled through a pair of secondary heat exchangers and passes through the condenser where excess water is extracted. The air then enters the ACM expansion turbine, where it is cooled to low temperature and reduced in pressure for discharge into the mixer. The heat exchangers are cooled by ram air, ducted into the aft fuselage by dorsal fin ram inlets (supplemented by the vent door to relieve negative pressure when the airplane is on the ground), and drawn into a ram-air duct within the air-conditioning bay by an ACM-driven fan. The duct directs the air through the primary and secondary heat exchangers before discharging it overboard through the tail cone. Water piped from the condenser is sprayed into the ram-air duct to improve heat exchanger efficiency. Cold air discharged into the mixer is blended with cabin air forced into the mixer by the cabin recirculation system and hot bleed air injected upstream of the mixer via an ACM bypass duct. By varying the blended ratios of cold air to hot bypassed air, the desired conditioned air temperature is achieved. Pack temperature control valves, one positioned in the ACM supply duct and one in the ACM bypass duct, regulate the blending ratios in response to commands from the cabin temperature controller.

Revision 2

21-25

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

POSITION MICROSWITCH RANGE CABIN VALVE

LINKAGE FLIGHT COMP VALVE

OPEN 90° 65° OPEN CLOSED 0°

MOUNTING PLATE

0° CLOSED

CCW STOP

CW STOP ACTUATOR ROTATION

FUNCTIONAL SCHEMATIC

FLIGHT COMPARTMENT CABIN BUTTERFLY VALVE BUTTERFLY VALVE

ACTUATOR

Figure 21-15. Temperature Trim Valves

21-26

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION AND OPERATION

NOTES

Temperature Trim Valve The temperature trim valve (Figure 21-15) consists of two one-inch butterfly valves driven by an electrically operated actuator through mechanical linkage. All three components are installed on a mounting plate attached to the condenser/mixer in the aft fuselage, aft of the rear pressure dome. The valve regulates the temperature of the air supplied to the flight compartment by opening and closing the butterfly valves to control the flow of hot bleed air to achieve the desired temperature. The valves are controlled by an electrically operated actuator which modulates in response to automatic signals from the flight compartment temperature controller, or from direct electrical inputs from the manually operated switch. The mechanical linkage connecting the two butterfly valves to the actuator is designed to provide a sequence of valve openings and closings, in response to actuator rotation, to provide a complete temperature range for the flight compartment. If a warmer flight compartment temperature is desired, the flight compartment butterfly valve is opened to allow hot bleed air directly into the distribution system to achieve the temperature. The cabin butterfly remains closed. If a cooler flight compartment temperature is desired, the cabin butterfly valve is opened to allow hot bleed air directly into the cabin distribution system. The flight compartment butterfly valve remains closed. Since the cabin has not called for this extra heat, the pack temperature control valve is repositioned to allow less bypass. The resulting cooler pack discharge temperature restores cabin requirements (pack discharge plus added heat) and gives the flight compartment its cooler temperature.

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COLD FULL-TRAVEL MICROSWITCH

2.5-INCH THROTTLE VALVE

PACK BYPASS BUTTERFLY VALVE

OPEN

VALVE POSITION

1.5-INCH BYPASS VALVE

CLOSED

CLOSED

CCW STOP

THROTTLE BUTTERFLY VALVE ACTUATOR

MOUNTING PLATE

CW STOP ACTUATOR STROKE FUNCTIONAL SCHEMATIC

LINKAGE

Figure 21-16. Pack Temperature Control Valves

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Pack Temperature Control Valves

NOTES

The electrically actuated temperature control valve consists of a 1.5-inch diameter pack bypass butterfly valve operated in opposition to a 2.5-inch diameter throttle butterfly valve (Figure 21-16). The two valves are mechanically linked to the actuator, and all three components are installed on a mounting plate attached to the heat exchanger in the rear fuselage aft of the rear pressure dome. The pack bypass valve is located in a bypass duct from the bleed-air duct; the throttle valve is in the outlet duct from the primary heat exchanger. The valve regulates the temperature of the air discharged from the pack by opening and closing of the pack bypass and throttle valves. The electrically operated actuator modulates in response to automatic signals from the cabin temperature controller or direct electrical inputs from the manually operated switch. The mechanical linkage connecting the two valves to the actuator provides a sequence of valve opening and closing, in response to rotation of the actuator, providing a complete temperature range from maximum cooling to maximum heating. For maximum cooling, the pack bypass valve is closed and the throttle valve is open. This allows the total flow from the engine bleed-air system to pass through the dual heat exchanger and air cycle machine. When less cooling is required, the actuator moves the throttle valve toward the closed position and the bypass valve toward open to allow a proportional amount of the airflow to bypass the refrigeration circuit. At maximum heat, the throttle valve is fully closed, and the pack bypass valve is fully open.

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RAM-AIR OUTLET HEADER

VENT DOOR

RAM-AIR OUTLET

BLEED BYPASS TO MIXING BOX INLET DUCT (FROM ENGINE BLEED OR APU [SOO 8062]) PACK TEMPERATURE CONTROL VALVE MOUNTING BRACKET

PRIMARY

SECONDARY

OUTLET DUCT (TO ACM COMPRESSOR) RAM-AIR INLET INLET DUCT (FROM ACM COMPRESSOR) OUTLET DUCT (TO ACM TURBINE)

TRANSITION DUCT RAM-AIR INLET HEADER

WATER NOZZLE

Figure 21-17. Heat Exchanger

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Heat Exchanger

NOTES

General The heat exchanger (Figure 21-17), located in the aft fuselage, is an all-welded aluminum alloy structure consisting of a housing with a finned core divided into primary and secondary units. Inlet and outlet headers on each side of the housing separate primary and secondary airflow. A circular transition duct, welded to a ram-air inlet header, functions as the ram-air inlet for the exchanger core. This duct is connected directly to the fan outlet end of the ACM. The primary exchanger inlet duct is connected to the bleed-air supply, and the outlet duct to the ACM compressor inlet. The secondary exchanger inlet duct is connected to the ACM compressor outlet and the outlet duct of the ACM turbine inlet through the condenser. Inward opening vent doors in the ram-air outlet header provide a means of pressure relief when the pressure differential in the air-conditioning bay reaches 0.4 psi. A spray nozzle in the ram-air inlet is connected to a water drain in the condenser and to another drain downstream of the condenser. Air tapped from the supply at the drains forces collected water to the nozzle to be sprayed into the heat exchanger inlet.

Heat Exchanger Operation The purpose of the heat exchanger is to lower the bleed-air temperature for the air cycle machine. Hot air from the bleed-air system flows to the primary heat exchanger. The airflow is partially cooled by heat transfer to ram air through the exchanger core and is then directed to the ACM compressor inlet. From the compressor, the air flows through the secondary heat exchanger where it is further cooled. The airflow is then directed through the condenser to and through the ACM turbine.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONDENSER MIXER

OUTLET TO CABIN AND FLIGHT COMPARTMENT

INLET FROM ACM TURBINE WATER DRAIN

INLET FROM RECIRCULATING FAN

CONDENSER OUTLET (TO ACM TURBINE)

WATER DRAIN

CONDENSER INLET (FROM HEAT EXCHANGER)

CONDENSER OUTLET

TECH C

Figure 21-18. Condenser/Mixer

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Condenser/Mixer

NOTES

General The condenser/mixer (Figure 21-18) is an allwelded aluminum alloy unit consisting of a finned core and inlet and outlet headers. It is bolted to structure in the aft fuselage. One end of the unit is the “mixer,’’ where cool air from the ACM, previously injected hot bypass bleed air, and recirculated cabin air mix to achieve the temperature required in the cabin. The other end of the unit is the condenser, where moisture is extracted from the air. Two drains collect extracted moisture and route it back to the heat exchanger inlet through a filter or the overboard drain.

Condenser/Mixer Operation Relatively cool conditioned air destined for the cabin passes through the finned core of the condenser. Warm air from the secondary heat exchanger passes over the cool core, where moisture in the air collects on the core. When droplets collect in water drains, they are routed back to the heat exchanger ram-air inlet or overboard drain.

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TURBINE OUTLET (TO CONDENSER/ MIXER) FAN INLET SCREEN

HEAT EXCHANGER CONNECTING FLANGE

TURBINE HOUSING

COMPRESSOR INLET (FROM PRIMARY HEAT EXCHANGER)

COMPRESSOR HOUSING

FAN OUTLET DUCT

FAN HOUSING

OIL LEVEL OIL FILL SIGHT GAGE COMPRESSOR OUTLET (TO SECONDARY PLUG HEAT EXCHANGER)

Figure 21-19. Air Cycle Machine

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Air Cycle Machine General The air cycle machine (ACM) (Figure 21-19) consists of a three-wheel assembly and separate housings for compressor, turbine, and fan rotors. The three housings are secured together at circular bolted flanges. The three-wheel assembly consists of a common shaft mounting a centrifugal compressor and turbine rotor on one end and an axial flow fan at the opposite end. The shaft rotates in a single bearing cartridge. The shaft seals are located to the front and rear of the bearing races to prevent oil leakage into the airflow. The bearings and the shaft are lubricated by oil from a sump in the turbine housing. Two wicks are radially preloaded against the shaft. The sump has drain and filler plugs and a sight gage. The dome-shaped compressor housing is flange-mounted on the turbine housing and contains a diffuser to direct the airflow leaving the centrifugal compressor. The compressor inlet is at the center of the domed housing and incorporates offset radial vanes to impart a swirling motion to the incoming air from the primary heat exchanger, thus correcting the impingement angle on the compressor rotor.

fan rotor. The fan rotates in a steel shroud, which is an integral part of the inlet. The outlet section is cone-shaped with a tubular bore and air-straightening vanes that provide a divergent transition duct for ram-air delivery to the heat exchangers. The diffuser outlet mates with the ram-air inlet of the heat exchanger.

ACM Operation Cooled air from the primary heat exchanger enters the compressor inlet of the ACM, where it is compressed and then delivered at a higher pressure and temperature to the secondary heat exchanger through the condenser to the turbine inlet. Expansion of the air across the turbine rotor reduces the pressure, with a corresponding drop in temperature. Air leaving the turbine outlet is routed to the condenser/mixer. The turbine extracts energy from the airflow as it reduces the pressure to just above cabin pressure. Most of this energy is fed back to aid in driving the compressor. The remainder of the turbine energy drives the fan to ensure airflow through the heat exchanger when the airplane is stationary.

NOTES

The turbine housing is the main structural member of the ACM, supporting the compressor and fan housings on the end flanges and the bearings and three-wheel assembly in the bore. A turbine nozzle in the housing aids in the expansion of air leaving the turbine rotor, lowering the air temperature. Tube adapters on the housing connect the turbine inlet to the secondary heat exchanger and the turbine outlet to the condenser/mixer. The fan housing consists of two sections, one section serving as the air inlet and the other section providing the air outlet. The inlet section provides a circular, reverse-flow duct which directs ram air through the axial-flow fan rotor. A wraparound wire mesh screen prevents the ingress of foreign objects, thus protecting the

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AIR CYCLE MACHINE

Figure 21-20. Compressor Discharge Overtemperature Switch

Figure 21-21. Temperature Controllers

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Compressor Discharge Overtemperature Switch

NOTES

The compressor discharge overtemperature switch (Figure 21-20) is installed in the delivery duct from the ACM compressor. It consists of a normally open single-pole thermal switch with a bimetallic element. The switch closes when the compressor discharge temperature exceeds 405 ±10° F (207 ±5° C).

Overheat Condition If the temperature of the air discharged from the ACM compressor exceeds 207° C (450° F), the compressor discharge overtemperature switch closes to energize both HP bleed control relays closed with the following results: • The AIR COND PACK HOT caution light comes on. • Both HP bleed valves are deenergized closed. • Both pressure regulator and shutoff valves are energized closed. When the overheat condition clears, the switch opens to restore system operation.

Temperature Controller The temperature controller (Figure 21-21) for each system is located in the electrical equipment bay forward of the No. 1 relay panel. They consist of aluminum boxes housing electronic analog devices with printed circuit boards. A circuit in each controller is connected to its associated temperature control selector and duct and zone temperature sensors. Another circuit connects the applicable valve actuator and manual temperature control switch through relays. By comparing the input signals from the duct and zone sensors with the selection made, the controller completes a circuit to operate the applicable valve actuator. The actuator modulates the valves in response to the controller signals to supply conditioned air at the desired temperature.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NOTE: PRE 1990 INTERIOR

TEMPERATURESENSING BULB

DUCT TEMPERATURE SENSOR

DUCT OVERTEMPERATURE SWITCH

Figure 21-22. Temperature Sensors and Switches

Figure 21-23. AIR CONDITIONING Control Panel

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Duct Temperature Sensors

Duct Overtemperature Switches

The duct temperature sensors (Figure 21-22) are negative coefficient thermistors in which the resistance varies inversely with the temperature of the air flowing through the duct. Each sensor provides a signal to the associated controller for automatic temperature control.

Each duct overtemperature switch consists of a normally open, single-pole thermal switch with a bimetallic element. When duct temperature exceeds 190° F (88° C), the switch closes to illuminate the appropriate cabin or flight compartment DUCT HOT caution light. Actuation of either switch also energizes a duct overheat control relay to switch the input to the pack temperature control valve actuator to the manual COOL command. The switch opens to resume air-conditioning system operation when the duct air temperature falls to 180° F (82° C).

NOTE 1990 interior temperature sensors and switches are relocated below the baggage compartment floor in the applicable ducts. The cabin duct temperature sensor is located in the cabin supply duct above the baggage compartment. The flight compartment duct sensor is in the flight compartment supply duct beneath the baggage compartment floor.

Zone Temperature Sensors The zone temperature sensors are similar in construction and operation to the duct sensors. They provide signals for automatic temperature control to the associated controller. The cabin zone temperature sensor with integral fan is located overhead in the center of the cabin area, left side. The flight compartment zone sensor is located behind the pilot in an exhaust duct into the avionics bay.

Temperature-Sensing Bulb The temperature-sensing bulb is an electrical resistance unit that transmits resistance va r i a t i o n s p r o p o r t i o n a l t o t e m p e r a t u r e changes to the cabin DUCT temperature indicator (Figure 21-23). The indicator reading allows for more precise temperature control when in the MAN mode.

Revision 2

OPERATION Air to be cooled is temperature-reduced through the primary heat exchanger and is then compressed through the ACM to boost the temperature and pressure (Figure 21-24). The compressed air passes through the secondary heat exchanger to further reduce its temperature, passes through the condenser where moisture is extracted, and then enters the ACM turbine. The energy extracted by the turbine drops the air temperature well below ambient and reduces the pressure to just above cabin ambient. At this point, bypassed bleed air is injected to achieve the desired temperature. The air then enters the mixing box, where it is blended with recirculated cabin air. Temperature of the air supply is regulated by opening and closing the pack temperature control valves to allow part of the hot bleed air to bypass the ACM and mix with the cool air discharged from the ACM. The valves are operated through mechanical linkage from a single actuator which positions the valves in response to signals from an automatic temperature controller (which can be monitored on the flight attendant’s panel) or from direct electrical inputs from a manually operated switch in the flight compartment.

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FLT COMPT DUCT HOT CABIN DUCT HOT

CABIN TEMPERATURE BIAS

HOT BLEED AIR

5

4 6

1

10 13

7

BAGGAGE

2

FLT COMP

12

11

8 3 9 CABIN

1. CABIN TEMPERATURE CONTROLLER 2. FLIGHT COMPARTMENT TEMPERATURE CONTROLLER 3. FLIGHT COMPARTMENT ZONE TEMPERATURE SENSOR 4. TEMPERATURE-SENSING BULB 5. CABIN DUCT OVERTEMPERATURE SWITCH 6. CABIN DUCT TEMPERATURE SENSOR 7. CABIN ZONE TEMPERATURE SENSOR (PRE MOD 8/0145)

8. FLIGHT COMPARTMENT TEMPERATURE SENSOR 9. FLIGHT COMPARTMENT DUCT OVERTEMPERATURE SWITCH 10. TEMPERATURE TRIM VALVE ACTUATOR 11. PACK TEMPERATURE CONTROL VALVE ACTUATOR 12. CONDENSER/MIXER 13. CABIN ZONE TEMPERATURE SENSOR (MOD 8/0145)

Figure 21-24. Temperature Control System

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TEMPERATURE CONTROL SYSTEM GENERAL Air-conditioning system temperature is controlled by two electrical subsystems, one for the cabin and the other for the flight compartment. The cabin subsystem controls temperature generated within the pack for the entire airplane; the flight compartment subsystem allows for changes in temperature bet w e e n t h e t wo a r e a s . E a c h s u b s y s t e m i s controlled by an independent set of switches on the AIR CONDITIONING panel, which can be positioned for manual or automatic operation.

adjustment of cabin temperature can be controlled by a five-position selector switch located at the forward flight attendant’s station. Each setting change provides a 2° F temperature adjustment in the direction selected (flight attendant’s temperature control panel, pre1990 interior).

NOTES

The cabin temperature control system utilizes a temperature controller that compares input signals from two sensors with the temperature selected and sends a signal to the pack temperature control valve actuator to modulate the control valves to achieve the desired temperature. The flight compartment temperature control system operates the same as the cabin system except the signal from the temperature controller is sent to the trim valve actuator to modulate the trim valves for the desired flight compartment temperature. Air-conditioning system controls on the overhead panel include a cabin DUCT temperature indicator and switches for temperature control (Figure 21-24). Temperature in the cabin and the flight compartment can be controlled either automatically or manually with two switches labeled “MAN’’ and “AUTO.’’ The two manual temperature controls are momentary-on springloaded to center switches sharing a COOL–WARM legend with the automatic control CABIN and FLT COMP rotary variable resistors. The TEMP CONTROL MAN–AUTO switches are wired into the electrical circuits of the appropriate temperature controller. Fine

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 21-25. Cabin Temperature Bias Control

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OPERATION

Flight Compartment

Automatic Temperature Control

Flight compartment air is initially at the same temperature as the cabin but can be adjusted by setting the FLT COMP automatic temperature control selector as required. This selection, plus the temperature sensed by the flight compartment duct temperature sensor in the aft section of the airplane below the baggage compartment floor and the flight compartment zone temperature sensor behind the pilot’s seat, forms a bridge circuit in the flight compartment temperature controller. The controller compares the two sensed temperatures with the selected temperature, and the difference, in the form of electrical impulses, is transmitted to the temperature trim valve actuator in the supply duct.

Cabin In the automatic mode, the cabin MAN–AUTO switch (Figure 21-23) is positioned to AUTO, and the CABIN automatic temperature control is rotated toward COOL or WARM, as required. Temperature of the cabin air is sensed by the cabin air duct temperature sensor in the inlet duct. Temperature of the air leaving the cabin is sensed by the cabin zone temperature sensor at the center left side of the cabin. The cabin zone sensor has an automatically powered fan to draw ambient cabin air over the sensor for a more accurate reading. These sensors are wired to form part of an electrical bridge circuit within the cabin temperature controller. An electrical input from the CABIN automatic temperature control selector, through the CABIN TEMPERATURE bias control, provides a variable in another leg of the bridge circuit. The controller compares the two sensed temperatures with the selected temperature, and the difference, in the form of electrical pulses, is transmitted to the pack temperature control valve actuator in the pack. The actuator responds to the signal by positioning the pack temperature control valve to regulate the flow of air at the desired temperature. Thus, the command output from the cabin temperature controller to the temperature control valve actuator brings the temperature at the zone sensor in the cabin in line with the selected temperature. The duct temperature sensor limits the air temperature to 35 to 160° F (2 to 71° C).

If a warmer temperature than that in the cabin is desired in the flight compartment, the actuator positions a trim valve to meter hot air into the flight compartment supply duct. If a cooler temperature is desired, the flight compartment trim valve closes, and the cabin trim valve opens to route hot air into the cabin supply duct. If additional heat has not been selected, the cabin system causes the pack temperature control valve to reposition to restore cabin temperature, with the added hot air, and to provide the desired temperature in the flight compartment.

NOTES

Cabin Temperature Bias Control The CABIN TEMPERATURE bias control (Figure 21-25), located at the flight attendant’s forward station, is used for minor adjustments in the desired cabin temperature. Connected in series with the CABIN selector on the AIR CONDITIONING panel, it modifies the signal sent to the cabin temperature controller to increase or decrease the cabin temperature by 2 to 4° F.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Manual Temperature Control Cabin In the manual mode, the cabin MAN–AUTO switch is positioned at MAN, and the required temperature is obtained by adjustment of the CABIN manual temperature control selector between COOL and WARM. In this mode, there are no limits in the supply temperature range. MAN selection also energizes a relay to initiate the following: •

Isolate the hot and cold command signals from the cabin temperature controller to the pack temperature control valve actuator.

•

Complete the cool and warm circuits between the CABIN manual temperature control selector and the pack temperature control valve actuator.

•

With the circuits complete, signals are sent to the actuator to position the control valves as determined by the selection made at the CABIN manual temperature selector.

If a warmer flight compartment temperature is desired, the trim valves are positioned to add heat to the flight compartment supply duct. If cooler temperature is desired, the trim valves are positioned to add heat to the cabin supply duct. The cabin will readjust to the added heat in the AUTO mode. If the CABIN manual selection is made, the trim valves lock in the center position if the trim valve attempts to add heat to the cabin system. To release the lock, the cabin must be selected to AUTO, removing the ground from the locking relay.

NOTES

Flight Compartment In the manual mode, the MAN–AUTO switch is positioned at MAN, and the required temperature is obtained by adjusting the FLT COMP manual temperature control selector between COOL and WARM. Selection of MAN also energizes a relay to initiate the following: •

Isolate the hot and cold command signals from the flight compartment temperature controller to the trim valve actuator in the supply duct.

•

Complete the cool and warm circuits between the FLT COMP manual temperature control selector and the trim valve actuator.

•

With the circuits complete, signals are sent to the actuator to position the valves as determined by the selection made at the FLT COMP selector.

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STAGE THREE

Duct Overheat If the cabin or flight compartment supply duct air temperature exceeds 190° F (88° C), a duct overtemperature switch closes with the following results: •

A circuit is completed to the master caution control box to illuminate a DUCT HOT caution light.

•

The duct overheat relay K3 or K4 (MSM Chapter 21) is energized with the following results: •

Relay K1 or K2 is energized to isolate the automatic hot or cold commands from either temperature controller to the associated valve actuator.

•

A circuit is completed to provide a cool signal to the associated valve actuator, causing the actuator to reposition its valves to operate the pack in the fully cold mode.

1. Pack valve goes full cold. 2. Microswitch causes return to center relay, driving trim valve to locked center position. When the duct temperature falls to 180° F (82° C), the overtemperature switch opens to resume air-conditioning system operation.

NOTES

Three stages of flight compartment overheat operation are as follows: STAGE ONE 1. Master CAUTION light illuminates. 2. Trim valve is disconnected for automatic and manual control. STAGE TWO Cabin in auto

Cabin in MANUAL

1. Trim valve adds heat to cabin ducts center.

1. Trim valve locks at center position.

2. Auto box senses heat and drives pack towards cold.

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DUCT OVERTEMPERATURE SWITCH

LAVATORY

DUCT TEMPERATURE SENSOR

OUTLET GRILLS

GASPERS

REAR PRESSURE DOME

CABIN DUCT TEMP SENSOR

FROM AIRCONDITIONING PACK

OUTLET GRILLS

FROM RECIRCULATION SYSTEM

ZONE TEMPERATURE TO FLIGHT SENSOR COMPARTMENT

FLIGHT COMPARTMENT BULKHEAD

Figure 21-26. Cabin Air Distribution LOWER CABIN OUTLET DUCTS

CABIN OVERHEAD DUCTING

DIVERTER VALVES

LOWER CABIN OUTLET DUCTS

CABIN UNDERFLOOR DUCTING

Figure 21-27. Diverter System

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CONDITIONED AIR DISTRIBUTION (PRE-1990 INTERIOR) SERIES 100

The duct from the pack incorporates a duct temperature sensor and an overtemperature switch, which are part of the temperature control system. A cabin duct temperature sensor provides signals to the cabin DUCT temperature indicator on the AIR CONDITIONING panel.

GENERAL

Diverter System

Conditioned air from the pack is ducted into the fuselage through the rear pressure bulkhead. A baffle directs 70% of the air to the cabin and 30% to the flight compartment. The duct divides forward of the bulkhead, directing air through ducts above the cargo compartment to the cabin outlet grilles. Flight compartment air is routed below the cabin floor. Conditioned air to the flight compartment is also used for side window demisting.

On airplanes with Mod 8/0192, outlet grilles are installed in the sidewall panels at floor level to give improved cabin heating during cold weather (Figure 21-27). A diverter valve in the cabin air duct above the baggage compartment redirects air from the upper duct to ducts under the cabin floor and then through flexible ducts to the floor level outlet grilles.

Additional airflow to controllable outlets (gaspers) in the cabin and flight compartment is provided by an air recirculation system. The fan in this system also assists in avionics compartment cooling.

CABIN AIR DISTRIBUTION The supply of conditioned air for the cabin enters at the center of the rear pressure dome (Figure 21-26). It is routed above the cabin roof to outlet grilles on both sides of the cabin roof and galley area.

Diverter valve operation is controlled by the CABIN AIR SUPPLY switch on the flight attendant’s panel (Figure 21-25). Selecting UPPER DUCT or LOWER DUCT causes the diverter valve to redirect conditioned air to the applicable outlet grilles. A light in the diverter valve switch illuminates when the motor has rotated the valve to either full travel position (full travel time is approximately 28 seconds).

NOTES

In addition to normal cabin air conditioning, manually controlled outlets (gaspers) are provided at each seat location and in the lavatory and flight compartment. The air for the gaspers is provided by a recirculation fan. Air is exhausted from the cabin through louvers into the avionics and baggage compartments. Air flowing into the avionics compartment for equipment cooling is picked up by the recirculation fan and recirculated to the gaspers and the condenser/mixer. Air flowing into the baggage compartment is vented overboard through the pressurization outflow valve.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SIDE WINDOW DEMIST OUTLETS

CABIN REAR PRESSURE DOME

DIVERTER VALVE (MOD 8/0563)

FLIGHT COMPARTMENT FAN FIXED OUTLETS IN FOOT WELL

FLIGHT COMPARTMENT BULKHEAD CABIN SUPPLY

ADJUSTABLE OUTLETS IN SIDE CONSOLE

DUCT TEMPERATURE SENSOR

SILENCER (MOD 8/0188) FLOW ADJUST LEVERS

DUCT OVERTEMPERATURE SWITCH

ZONE TEMPERATURE SENSOR

SIDE WINDOW DEMIST OUTLETS FWD PRESSURE BULKHEAD

Figure 21-28. Flight Compartment Air Distribution

DIVERTER VALVE REAR PRESSURE DOME

CONS

AIR DIVERT FLT COMP HOLD

TEE DUCT ASSEMBLY

CABIN

SWITCH

5

5

TAXI PWR

TAXI CONT

LOG CONT 1

L FLARE

AIR STAIR

20

5

5

25

5

EXT LIGHTS

CABIN LTS PWR 1 CONT 15

5

A

B

LOGO L

LOGO R

C

D

E

LEFT DC CIRCUIT-BREAKER PANEL

Figure 21-29. Diverter Valve Installation (Mod 8/0563)

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FLIGHT COMPARTMENT AIR DISTRIBUTION

NOTES

The supply of conditioned air for the flight compartment, consisting of approximately 30% of the air discharged from the air-conditioning pack, enters the cabin at the center of the rear pressure dome and is routed to the flight compartment via flexible ducts underneath the cabin floor (Figure 21-28). On airplanes with Mod 8/0563, additional air is supplied to the flight compartment by a diverter valve (Figure 21-29) located in the main supply duct forward of the rear pressure dome. A selector switch on the left DC circuit-breaker panel, when activated, increases airflow to the flight compartment to 60%. A wye duct splits the airflow and directs it into two individual but identical air distribution subsystems, one for the pilot and one for the copilot. A fan provides air for flight compartment ventilation when the airplane is on the ground. On airplanes with Mod 8/0188, a silencer is installed in the flight compartment air supply duct to reduce air noise. Each of the two air distribution subsystems consists of a side window demist twin outlet, a fixed outlet in the foot well, and three adjustable outlets in the side console. Flow adjustment levers on the sill below the side window are connected to push-pull controls which operate movable vanes in the supply ducts to permit adjustment of airflow. A manually controllable air outlet (gasper) is located on the pilot’s and copilot’s side panels. The air for the gaspers is provided by the recirculation fan. Air is exhausted from the flight compartment through louvers into the avionics compartment. The exhausted air is picked up by the recirculation fan, which recirculates the air to the gaspers and to the air-conditioning pack and ventilates the compartment for cooling purposes.

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RECIRCULATED AIR

SIDE WINDOW DEMIST FLIGHT COMP GND FAN GASPER 42 FLIGHT COMP ZONE TEMP SENSOR FLOW ADJUST RECIRCULATION FAN

FLIGHT COMP SUPPLY

FIXED GRILLE INBOARD OF FOOT WELL

THREE ADJUSTABLE GRILLES IN THE SIDE CONSOLE

AIR SUPPLY GRILLES

Figure 21-30. Flight Compartment and Fan System FC FAN PWR

A3 A2

10 28 VDC R SEC BUS FC FAN CONT 5 28 VDC R SEC BUS

A1 OFF FC FAN

A B

K1

FAN CONTROL RELAY

PROCXIMITY SWITCH ELECTRONIC UNIT

LGWOW1 RGWOW1 GROUND PROVIDED WITH WEIGHT ON BOTH WHEELS (WOW)

Figure 21-31. Flight Compartment Fan Electrical Schematic

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Flight Compartment Fan

NOTES

The flight compartment fan picks up air from the nose wheel well area and routes it to the two flight compartment air distribution subsystems (Figure 21-30). Two check valves in the ducting prevent airflow from the fan from entering the flight compartment supply duct. Check valves in the ducting prevent conditioned air from escaping into the nose area when the air-conditioning system is operating.

Fan Operation With weight on both main gears, selecting the FC FAN switch on the AIR CONDITIONING panel to on energizes the fan control relay K1 (Figure 21-31) to complete a circuit from the 28-VDC right secondary bus to the fan. With either main gear weight off the wheels, the relay deenergizes to open the fan circuit. The ground to the relay is applied or removed through the PSEU.

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Figure 21-32. Recirculation Fan REAR PRESSURE DOME

GASPERS

AVIONICS COMPARTMENT

FILTER

TO PACK MIXING BOX

RECIRCULATION FAN

CHECK VALVE

RECIRC FAN PWR 28 VDC LEFT SEC BUS

A

50A CR1

RECIRC FAN CONT 28 VDC 5A LEFT SEC BUS

OFF

K1 FAN CONTROL RELAY

B RECIRCULATION FAN

RECIRC

Figure 21-33. Air Recirculation System

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AVIONICS COMPARTMENT COOLING Avionics compartment cooling is accomplished by a recirculation fan located under the cabin floor (Figure 21-32). Air exhausted from the cabin and flight compartment enters the avionics compartment through louvers in the walls of the compartment. The air is drawn over the equipment by the recirculation fan and is then directed to the gaspers or the pack.

AIR RECIRCULATION SYSTEM General The air recirculation system provides recirculated cabin air to the pack and to individual controllable gaspers in the flight compartment and lavatory.

through louvers in the compartment walls. It is drawn over avionics equipment by the fan and then directed back to the gaspers and the condenser/mixer. Normally, the recirculation system operates continuously to ventilate the avionics compartment and to provide airflow to the cabin and flight compartment. The recirculation fans incorporate protective features for undervoltage, overcurrent, and overheat. The symptoms are those of fan failure. Recirculation fans automatically shut down when exposed to transient voltages of less than 22 VDC for periods in excess of 0.10 second. They automatically restart between 18 and 20 VDC. On occasion, an overcurrent condition occurs; the fan must be switched off for 20 seconds prior to restart. Overheat also causes the fan to shut down; it automatically restarts when the overheat condition clears.

NOTES

The system consists of a fan, a check valve, and associated ducting. The fan is controlled by the RECIRC switch on the AIR CONDITIONING panel. Air supply for the fan is drawn over the avionics equipment. A duct extends rearward beneath the cabin floor to the condenser/mixer, where recirculated cabin air is mixed with conditioned air. A check valve in the duct prevents backflow of conditioned air when the recirculation system is not in operation. Secondary ducts extend up the cabin walls into the ceiling to supply recirculated air to the gaspers in the cabin, flight compartment, and lavatory.

Operation Selecting the RECIRC switch on completes a circuit to energize fan control relay K1 and complete a circuit from the 28-VDC left secondary bus to the recirculation fan (Figure 21-33). Air exhausted from the cabin and flight compartment enters the avionics compartment

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21-54 DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AVIONICS COMPARTMENT COOLING

Canada Ltd.

Revision 2

Figure 21-34. Conditioned Air Distribution (1990 Interior)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONDITIONED AIR DISTRIBUTION (1990 INTERIOR)— SERIES 100A

NOTES

GENERAL Conditioned air from the pack is ducted into the fuselage at the center of the rear pressure dome bulkhead (Figure 21-34). A longitudinal baffle in the duct directs 70% of the air to the cabin and 30% to the flight compartment via two ducts. The flight compartment duct is routed under the baggage compartment and cabin, where it splits into two ducts to feed the pilot’s and copilot’s subsystems. The cabin supply duct is routed under the baggage compartment, where it splits into an upper and lower duct for each side of the fuselage. The upper of these ducts supplies air to the cabin dado panel grilles, while the lower duct supplies, via sidewall risers, the upper cabin air outlets, as well as the cabin, lavatory, and flight attendant’s station gaspers. The flight compartment duct supplies air to sidewall grilles and gaspers in the flight compartment. Flight compartment air is also used for side window demisting. On airplanes with SOO 8069, conditioned air can be supplied to the cabin and flight compartment through an eight-inch universal ground air service connector located on the right side of the aft fuselage. The air supply from the ground connection is routed through a short flexible duct to join the recirculated air duct in the aft fuselage. The air is then routed, via the condenser/mixer, to the cabin and flight compartment. A butterfly-type check valve is line-mounted in the ground conditioned air duct to prevent recirculated or air-conditioning pack air from spilling overboard through the ground air connector.

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21-56 UPPER GRILLE GALLERY

GASPER GALLERY

DADO PANEL

RISER DUCT OVERTEMPERATURE SWITCH

MAIN CABIN SUPPLY DUCT

DUCT TEMPERATURE SENSOR

DUCT TEMPERATURE SENSING BULB

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT ATTENDANT CONTROLLED SHUTOFF VALVE

RESTRICTOR

TO FLIGHT ATTENDANT AND LAVATORY GASPERS

REAR PRESSURE BULKHEAD

Canada Ltd.

Revision 2

Figure 21-35. Cabin Air Distribution Schematic (1990 Interior)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN AIR DISTRIBUTION The supply of conditioned air for the cabin, consisting of approximately 70% of the air discharged from the air-conditioning pack, enters the cabin at the center of the rear pressure dome (Figure 21-35). The air is routed below the cabin to dado panels above the cabin floor and, via sidewall risers, to the upper air outlets, which are located outboard of the passenger service (PSU) units. The supply duct splits at a wye duct to supply each side of the cabin. At the wye split, a secondary upper duct branches out from the lower duct and parallels it on each side of the fuselage. The upper duct supplies air to the cabin dado panels, while the lower duct supplies air to the upper air outlets and gaspers via sidewall risers. Incorporated into the upper duct is a motorized butterfly shutoff valve controlled by a two-position switch on the flight attendant’s panel. The switch positions are UPPER DUCT and LOWER DUCT. In the UPPER DUCT position, the shutoff valve shuts off air to the cabin dado panel, routes all air to the upper air outlets, and gives maximum flow to the gaspers. When LOWER DUCT is selected, the shutoff valve opens to allow half of the air to the floor dado panels.

temperature sensing bulb, located in the cabin supply duct downstream of the overtemperature switch, provides signals to a cabin duct temperature indicator on the AIR CONDITIONING panel in the flight compartment.

NOTES

In addition to normal cabin air conditioning, manually controllable air outlets (gaspers) are provided at each seat location, in the center ceiling above the forward flight attendant’s station, the lavatory, and the flight compartment. All gaspers, except those in the flight compartment, receive conditioned air from the cabin air supply duct via takeoff ducts from the same sidewall risers as the upper cabin (PSU) outlets. Approximately 60% of the cabin air is exhausted through slots located between the side and dado panels of the mid and aft cabin, passing under the cabin floor to be exhausted through the normal outflow valve. The other 40% of cabin air is exhausted through the air recirculation system. The delivery duct from the air-conditioning pack incorporates a duct temperature sensor and an overtemperature switch, which are part of the temperature control system. A cabin duct

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FORWARD PRESSURE BULKHEAD

REAR PRESSURE BULKHEAD

SIDE WINDOW DEMIST OUTLETS

FLOW ADJUST LEVERS CABIN

TO LARGE FORWARD SIDE CONSOLE GASPERS

FIXED OUTLETS IN FOOT WELL

SIDE CONSOLE ADJUSTABLE OUTLETS

DUCT OVERTEMPERATURE SWITCH DUCT TEMPERATURE SENSOR

TO GASPERS BELOW WINDSCREEN CORNER PILLARS

CABIN SUPPLY

FLOW ADJUST LEVERS

ZONE TEMPERATURE SENSOR SIDE WINDOW DEMIST OUTLETS

Figure 21-36. Flight Compartment Air Distribution (Mod 8/1538)

LAVATORY GASPER

FLIGHT ATTENDANT STATION GASPER

GASPER GALLERY

REAR PRESSURE BULKHEAD

FLIGHT COMPARTMENT

CABIN RESTRICTOR RISERS

LOWER CABIN AIR FLIGHT COMPARTMENT AIR LAVATORY

Figure 21-37. Gasper System Schematic (1990 Interior)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT AIR DISTRIBUTION The supply of conditioned air for the flight compartment, consisting of approximately 30% of the air discharged from the air-conditioning pack, enters the cabin at the center of the rear pressure dome and is routed to the flight compartment via a duct under the cabin floor (Figure 21-36). A wye duct splits the airflow into two individual but identical air distribution subsystems, one for the pilot and one for the copilot. Each of the subsystems consists of a side window demist twin outlet, a fixed outlet in the foot well, and an adjustable large outlet in the side console. Flow adjustment levers on the sill below the side window are cable connected to operate movable vanes in the supply ducts to permit adjustment of airflow. Both crewmembers are provided with one large and one small manually controllable (gasper) air outlet. The large gaspers are at the forward end of each side console. The small gaspers are located below the windshield pillars on both sides of the flight compartment.

to individual controllable outlets at each passenger seat. The risers also supply conditioned air to the cabin upper air outlet grilles. The upper cabin grilles require less airflow than the gaspers; therefore, a restrictor is installed in each riser between the gasper tapoff and upper grille gallery. An additional single riser supplies cabin duct air to a lavatory gasper and a flight attendant station ceiling gasper. The flight compartment gasper system is supplied with conditioned air from the flight compartment conditioned air supply duct. Two individual controllable outlets are provided for the pilot and the copilot. A large outlet is located at the forward end of each side console, and a smaller outlet is located below the windshield corner pillars.

NOTES

Air is exhausted from the flight compartment under the floor to the recirculation fan. From there the air is drawn back to the air-conditioning pack condenser/mixer by the recirculation fan. The delivery duct from the air-conditioning pack incorporates a duct temperature sensor and an overtemperature switch, which are part of the temperature control system.

GASPER AIR SYSTEM The gasper air system (Figure 21-37) comprises two separate and independent systems, one for the cabin and one for the flight compartment. The one for the left system is supplied with conditioned air from the lower cabin conditioned air supply duct tapoffs from six sidewall risers (three on each side of the fuselage). The risers route air to the gasper galleries that extend above the passenger service units (PSU) panels on each side of the cabin. From there the air is supplied

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPT

AVIONICS COMPT

OUTFLOW VALVE CABIN

AIR EXTRACTION SLOTS

BAGGAGE COMPT CHECK VALVE

SLOT BETWEEN SIDE AND DADO PANEL

RECIRCULATION FAN REAR PRESSURE BULKHEAD

RECIRCULATION DUCT RECIRC FAN PWR 28 VDC LEFT SEC BUS

A

50A CR1

RECIRC FAN CONT 28 VDC 5A LEFT SEC BUS

OFF

2123K1 FAN CONTROL RELAY

B RECIRCULATION FAN

RECIRC

Figure 21-38. Air Recirculation System Schematic (1990 Interior)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR RECIRCULATION SYSTEM The air recirculation system (Figure 21-38) supplies air to the condenser/mixer of the airconditioning pack, where it is blended with conditioned air from the air cycle machine. The system consists of a recirculation fan, check valve, and associated ducting. The fan is controlled, through a fan control relay, by a switch marked OFF–RECIRC on the airconditioning panel. Electrical power is provided by the 28-VDC left secondary bus. The fan draws air back through the recirculation duct from under the flight compartment floor and from the cabin via adjoining air extraction ducts, located behind the top of the forward cabin dado panels, to the air-conditioning pack condenser/ mixer. The fan is located aft of the rear pressure bulkhead adjacent to the pack. The check valve is installed between the fan and the pack to prevent back-flow of conditioned air when the recirculation system is not in operation.

NOTE The fans incorporate protective features for undervoltage, overcurrent, and overheat. The symptoms are those of a fan failure. The fans automatically shut down when exposed to transient voltages of less than 22-VDC for more than 0.10 second. They automatically restart between 18 and 20 VDC. On occasion, an overcurrent condition will occur; the fan must be switched off for 20 seconds prior to restart. Overheat will also cause shut down, but it will automatically restart when the overheat condition clears.

Approximately 40% of the cabin air is recirculated; the other 60% is exhausted through slots located between the dado and side panels of the mid- to rear-cabin area. The exhausted air passes through the slots to the underfloor area and is exhausted through the normal outflow valve on the rear pressure bulkhead.

Operation Selecting the control switch to RECIRC completes a circuit to energize fan control relay 2123-K1. The relay then completes a circuit from the 28-VDC left secondary bus to the recirculation fan.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

LOW SPEED WARNING DEVICE EQUIPMENT THERMOSTATIC COOLING FAN SWITCH AIR EXTRACTION DUCT

FLT COMPT

AVIONICS COMPT

CABIN

FLT COMPT SENSOR

AVIONICS FAN

P1

LOW SPEED WARNING DEVICE FEABC

(20 SEC DELAY ON DIODE OUTPUT PIN F) WARNING LIGHT DS1

C B A

A3 A1

FAN

B3 B2 A2

A2

J2 P2

B1 A3 A1

X1 X2

K2

X1 X2

K1

S1 CLOSES AT > 35°C

COOLING FAN

10 28V DC L MAIN BUS P/O LEFT AVIONICS CIRCUIT BREAKER PANEL

NOTE: IDENT CODE IS 2126, UNLESS OTHERWISE INDICATED

Figure 21-39. Avionics Compartment Cooling Fan

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AVIONICS COMPARTMENT COOLING (1990 INTERIOR) Avionics compartment cooling is accomplished by an avionics fan located under the cabin floor (Figure 21-39). The fan draws air off each of the equipment shelves, via air circulation ducts, and discharges it under the cabin floor. The fan is a brushless, thermostatically controlled unit that is operational when DC power is applied to the airplane. In the event of fan failure, cooling by natural convection is adequate for interim operation. Fan failure is indicated by a light on the forward face of the flight attendant’s panel.

NOTE A 20-second delay is incorporated in the low-speed warning device on output pin F. This prevents a false fan failure indication while the fan is spooling up to operational speed after initial start.

OPERATION Normal The avionics cooling fan operates when the ambient temperature at the avionics rack exceeds 35° C (as detected by the thermostatic switch mounted on the upper avionics shelf). In this condition, switch S1 closes and relay K1 is energized. 28 VDC is applied through the fan circuit breaker and contacts of relay K1 and K2 to energize the fan and provide power to the low-speed warning device. Normal operating speed for the fan is 9,000 rpm.

Fan Failure Speed of the fan is monitored at pin C of the lowspeed warning device by an output from the fan. If the speed of the fan drops to 6,000 rpm or less and remains at the reduced speed for more than 20 seconds, the low-speed warning device provides an output which energizes relay K2. Contacts of K2 disconnects power from the fan; K2 also maintains power to the low-speed warning device independent of relay K1. The low-speed warning device also turns on the diode warning lights DS1 to indicate fan failure.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NOSE GEAR WELL FLIGHT COMPT

CARGO COMPT

CABIN

UNDERFLOOR AREA AFT PRESSURE DOME

LEGEND FORWARD PRESSURE BULKHEAD

PRESSURIZED

UNPRESSURIZED

Figure 21-40. Pressurized Areas

3

2

6

CAB ALT

1 0 -1

4 5

1000 ft

7 8 9 10

BA RO in H G 31 30 29 28

ALT

BAR

CABIN ALTITUDE DUMP RATE F M A A U N L T AUTO CAB SET

INCR

NORM

Figure 21-41. Pressurization Controls and Indicators

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PRESSURIZATION SYSTEM—SERIES 100

NOTES

GENERAL NOTE The installation of two normal outflow valves on Series 300 airplanes is the only pressurization system difference between Series 100 and Series 300 airplanes. Pressurization of the airplane is dependent on three factors: (1) a positive, controlled airflow provided by the bleed-air system, (2) the flight compartment and cabin area being appropriately sealed, and (3) a controlled rate of air escaping from the fuselage. Air exhausted from the cabin and flight compartment is metered overboard through an outflow valve in the rear pressure dome. Valve operation depends on the setting of a pressure control unit on the overhead panel. A manually operated safety outflow valve in the forward pressure bulkhead can be used for a backup pressurization control or for smoke control. Pressurization is controlled by a normal outflow valve that modulates in response to electrical signals generated by a cabin pressure controller. The pressurized area of the fuselage is shown in Figure 21-40. The system is monitored from the selector and indicator panels (Figure 21-41). The selector panel provides for operation of the system in the automatic, manual, or dump modes. The indicator panel is used to monitor the system.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PRESSURE RELIEF (SET TO RELIEVE AT 5.8 PSI)

CABIN AMBIENT

SIGNAL FROM PRESSURE CONTROLLER

CAB STATIC CAB

TORQUE MOTOR 18 PSI FROM DEICE SYSTEM

VENTURI

Figure 21-42. Normal Outflow Valve—Series 100

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION AND OPERATION

NOTES

Normal Outflow Valve— Series 100 The normal outflow valve (Figure 21-42) consists of a poppet and diaphragm assembly, a spring, an outer dome, a torque motor, and a differential pressure limiter. The valve is pneumatically operated and springloaded closed. Cabin pressure is applied to the inside of the poppet/diaphragm assembly and to the outer dome. Selecting the desired cabin altitude generates a DC signal to position the torque motor to meter suction produced by the venturi ejector to the outer dome. Thus, a pressure differential is created between each side of the poppet and diaphragm assembly. If the cabin altitude is less than selected altitude, the valve opens to bleed some of the cabin pressure, increasing the cabin altitude. A drop in cabin pressure is sensed as a reduction of poppet opening force, and the outflow valve moves toward the closed position to restore cabin pressure (decrease cabin altitude). If pressure differential between the cabin and ambient exceeds 5.8 ±0.15 psi, a differential pressure limiter in the outflow valve opens to connect the outer dome to ambient, bleeding pressure from the dome and opening the valve. The valve also provides negative pressure relief and opens if outside ambient pressure exceeds cabin pressure. The safety outflow valve on the forward pressure bulkhead is similar in operation but is controlled manually with a manual control needle valve on the control unit. Smoke can be removed from the flight compartment by opening the valve.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

INDICATOR PANEL

CONTROL UNIT AND SELECTOR PANEL

MASTER NORMAL OUTFLOW CABIN PRESSURE VALVE CONTROLLER

FORWARD PRESSURE BULKHEAD

18 PSI FROM DEICE SYSTEM

EJECTOR SAFETY OUTFLOW VALVE

FORWARD DUMP MANUAL SELECTOR

SUCTION SLAVE NORMAL OUTFLOW VALVE

TRUE STATIC

REAR PRESSURE DOME

STATIC SUCTION PRESSURE RELIEF (SET TO RELIEVE AT 5.8 PSI)

FILTERED CABIN AMBIENT ELECTRICAL SIGNAL FROM CABIN PRESSURE CONTROLLER

POPPET/DIAPHRAGM ASSEMBLY OUTER DOME

TO MAN KNOB ON CONTROL PANEL

CAB STATIC CAB

CAB CAB

CAB

PRESSURE RELIEF (SET TO RELIEVE AT 5.8 PSI)

SUCTION EJECTOR

TORQUE MOTOR

XXXXX * CABIN AMBIENT

STATIC

* WHEN OUTER DOME PRESSURE EQUALS CABIN AMBIENT, THE VALVE IS CLOSED; APPLYING SUCTION OPENS THE VALVE.

PRESSURE RELIEF (SET TO RELIEVE AT 5.8 PSI)

PORT PLUGGED

CAB TORQUE MOTOR NOT ELECTRICALLY CONNECTED

STATIC CAB

PORT PLUGGED

Figure 21-43. Pressurization Control Schematic—Series 300

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Normal Outflow Valves—Series 300 Pressurization is controlled by two normal outflow valves (Figure 21-43) which modulate in response to electrical signals generated by the cabin pressure controller. The valves are located on the rear pressure dome. The right valve is the main (master) valve and is the only valve having the torque motor receiving electrical commands from the cabin pressure controller. The left valve is the auxiliary (slave) valve. It has a torque motor which is not connected to the cabin controller, but the valve is connected pneumatically to the master valve and instantly follows its positions as commanded by the cabin pressure controller. Both valves consist of a torque motor, poppet/diaphragm assembly, spring, outer dome, and differential limiter. The valves are pneumatically operated open and are spring-loaded closed. Cabin pressure is applied to the inside of the poppet/diaphragm and to the outer dome of each valve.

ambient air pressure by 5.8 ±0.15 psi. The maximum regulated differential pressure permitted by the cabin pressure controller is 5.5 ±0.3 psi, which corresponds to an 8,000 foot cabin at 25,000 feet.

Negative Pressure Relief Valves Protection from negative pressure is also provided by the normal and safety outflow valves, which automatically open if the interior suction exceeds 0.1 psi.

NOTES

Venturi Ejector The venturi ejector is attached to the pressure dome adjacent to the outflow valve. Air pressure at approximately 18 psi from the airframe deicing system is directed to the venturi to produce suction, which is then applied to the outer dome of the normal outflow valve through the torque motor.

Safety Outflow Valve The safety outflow valve is similar in construction to the normal outflow valve. The safety outflow valve employs suction generated by venting one side of the valve to the slipstream; the suction is modulated by the manual metering valve. Thus, the system requires no electrical power or pneumatic pressure to function when the airplane is in flight.

Pressure Relief Valves Both the safety and normal outflow valves contain pressure relief valves which vent excess pressure if cabin pressure exceeds sensed

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN PRESSURE RATE-OF-CHANGE SELECTOR KNOB CONTROLS CABIN ALTITUDE RATE OF CLIMB/DESCENT WHEN SYSTEM IS IN SEMIAUTOMATIC MODE. FAULT LIGHT ILLUMINATES TO INDICATE SYSTEM MALFUNCTION. ALSO ILLUMINATES BRIEFLY DURING SYSTEM SELF-TEST.

CABIN ALTITUDE SETTING INDICATOR POINTER INDICATES CABIN ALTITUDE REQUIREMENT SET BY ALT KNOB.

MANUAL CONTROL KNOB DIRECTLY MODULATES SAFETY OUTFLOW VALVE SUCTION. TURNING CLOCKWISE OPENS OUTFLOW VALVE, CAUSING CABIN ALTITUDE TO INCREASE. CONTROL SENSITIVITY IS REDUCED AT LOWER ALTITUDES.

CABIN ALTITUDE KNOB INPUTS CABIN ALTITUDE REQUIREMENT INTO CABIN PRESSURE CONTROLLER.

3

2

6

CAB ALT

1 0 -1

4 5

1000 ft

7 8 9 10

BA RO in H G 31 30 29 28

ALT

BAR

CABIN ALTITUDE DUMP RATE F M A A U N L T AUTO CAB SET

INCR

NORM

BAROMETRIC CORRECTION INDICATOR

MODE SELECTOR SWITCH

BUG SHOWS CORRECTION SET ON BAR KNOB.

LEVER LOCKED SWITCH HAS THREE POSITIONS: AUTO — CABIN PRESSURE CONTROLLER IS ACTIVATED TO OPERATE FULLY AUTOMATICALLY OR SEMIAUTOMATICALLY, DEPENDING IN POSITION OF AUTOMATIC FUNCTION SWITCH.

BAROMETRIC CORRECTION KNOB INPUTS DESIRED BAROMETRIC CORRECTION INTO CABIN PRESSURE CONTROLLER AUTOMATIC FUNCTION SWITCH WITH MODE SELECTOR AT AUTO: NORM —SELECTS FULLY AUTOMATIC OPERATION CAB SET —SELECTS SEMIAUTOMATIC OPERATION

MAN —

CABIN PRESSURE CONTROLLER IS DEACTIVATED AND PRESSURIZATION CONTROL IS THROUGH SAFETY OUTFLOW VALVE VIA MANUAL CONTROL KNOB.

DUMP — NORMAL OUTFLOW VALVE IS HELD OPEN TO PREVENT AIRCRAFT FROM PRESSURIZING.

Figure 21-44. Selector Panel

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Selector Panel The selector panel (Figure 21-44) contains a combined cabin altitude and barometric indicator with ALT, BARO, and RATE knobs, an AUTO–MAN–DUMP switch, a CAB SET–NORM switch, a control needle valve marked “INCR,’’ and a FAULT light. The BARO knob is provided to set barometric pressure, the desired cabin cruising altitude/destination altitude is set with the ALT knob, and cabin altitude rate of change is set with the RATE knob. The RATE knob is rotated clockwise to increase the rate of change. When the ball on the knob is aligned with the index mark on the controller, the cabin rate of change is 500 fpm up and 300 fpm down. Movement full counterclockwise sets a minimum fpm and full clockwise sets the maximum rate of 0 to 2,500 fpm up and 1,500 fpm down.

The control needle valve is mechanically connected to the outer dome of the front safety outflow valve and to ambient. An arrow on the panel indicates that a clockwise selection opens the needle valve, venting the outer dome and increasing the cabin altitude. Full counterclockwise rotation closes the needle valve. The FAULT light provides indication of any fault in the pressurization control system. It illuminates for approximately two seconds when electrical power is first applied, indicating that dynamic self-testing is in progress. The light goes out if no fault is detected during the selftest, and passive continuous self-testing continues.

NOTES

The three-position AUTO–MAN–DUMP switch selects the mode of operation for pressurization control. In the AUTO position, control is completely automatic or semiautomatic, depending on the CAB SET–NORM switch position. In the MAN position, the desired cabin pressure is controlled by rotating the control needle valve toward INCR to increase cabin altitude or counterclockwise to decrease cabin altitude. In the DUMP position, the normal outflow valve is commanded to a fully open position and the airplane may be operated unpressurized. The CAB SET–NORM switch allows selection of pressurization in the CAB SET position or the cabin pressure controller to control pressurization automatically in the NORM position. In the CAB SET position, the crew selects the desired cabin cruising altitude and then the destination altitude (as descent begins) with the ALT knob. The cabin altitude rate of change can be varied with the RATE knob. In the NORM position, the destination altitude is selected prior to takeoff, and pressurization is automatic from takeoff to touchdown. In both cases, the BARO knob must be set to the correct barometric pressure.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN RATE OF CLIMB INDICATOR REGISTERS CABIN INTERIOR PRESSURE CHANGES WHICH ARE INDICATED AS A RATE OF CLIMB OR DESCENT

CABIN F OF

OFF

0

1

6

2

5 3

4

DIFF PSI

CABIN PRESS WARNING LIGHT ILLUMINATES TO INDICATE CABIN PRESSURE EXCEEDS 10,000 FEET.

30

0 2

20 14 12 10

4 6 8

FTx ALT 1000

1

2

UP

0 DOWN

1 RATE

2 FPMx 1000

DIFFERENTIAL PRESSURE INDICATOR

CABIN ALTITUDE INDICATOR

REGISTERS DIFFERENCES IN INTERIOR AND EXTERIOR PRESSURE (NORMAL MAXIMUM 5.5 ± .3 PSI). POINTER RESTING AT OFF INDICATES NO POWER TO INDICATOR.

REGISTERS CABIN PRESSURE IN TERMS OF EQUIVALENT ALTITUDE. POINTER RESTING AT OFF INDICATES NO POWER TO INDICATOR.

Figure 21-45. Indicator Panel Table 21-1. PRESSURIZATION CONTROL SETTINGS SELECTOR PANEL SETTINGS AUTO/MAN/DUMP

CAB SET/NORM

RATE

ALT

BAR

MAN

AUTO

NORM

(INDEX MARK)

SET TO DESTINATION ALTITUDE

SET ATMOSPHERIC PRESSURE

FULL COUNTERCLOCKWISE

Electrical output from the computer controls suction applied to the normal outflow valve. Cabin altitude is maintained at, or as close as possible to, destination altitude without exceeding rate or differential pressure limits. AUTO

CAB SET

AS DESIRED

ENROUTE THEN DESTINATION

SET ATMOSPHERIC PRESSURE

FULL COUNTERCLOCKWISE

Electrical output from the computer controls suction applied to the normal outflow valve. Cabin altitude approaches and maintains enroute altitude. Destination altitude must be set at start of descent. MAN

N/A

N/A

N/A

N/A

MODULATE

Normal outflow valve is not used and is closed. Rotating manual dial meters low ambient pressure to modulate the safety outflow valve. Control effectiveness decreases as cabin pressure is reduced toward ambient. DUMP

N/A

N/A

N/A

N/A

N/A

Electrical output from the computer increases to apply full section to hold the normal outflow valve fully open.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Cabin Pressure Computer

NOTES

The cabin pressure computer, located on the avionics rack, controls the operation of the normal outflow valve. It compares dialed-in signals from the selector panel with cabin pressure, ambient pressure, and its computer schedule in relation to given information. It then sends an electrical signal to open or close the outflow valve to maintain the selected cabin pressure.

Indicator Panel The indicator panel (Figure 21-45) consists of a DIFF pressure indicator, a cabin ALT indicator, and a cabin altitude RATE-of-change indicator. The indicators, receiving signals from the ADC, are used to monitor the pressurization control system in both the AUTO and MAN modes.

OPERATION General The pressurization control system operation is primarily electrical. The pressurized area of the fuselage is supplied with a relatively constant flow of conditioned air from the bleed-air systems. Pressure in the fuselage is maintained by modulating the normal outflow valve to regulate the amount of air discharged to ambient. Both the normal and the safety outflow valves incorporate integral differential pressure limiters to open the valves when differential pressure exceeds a predetermined value. Both valves also open for negative pressure relief if the outside ambient pressure exceeds fuselage pressure. A summary of control settings for the various modes of operation is shown in Table 21-1.

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21-74

LEGEND B

DEICER PRESSURE CONTROL PRESSURE

AIR DATA COMPUTER

A CABIN PRESSURE CABIN AIR ATMOSPHERIC PRESSURE

NORM OPEN

ELECTRICAL

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

P VALVE 5.8 PSI

CABIN ALTITUDE DUMP

CABIN F OF

F OF 0

0 1

6

30 2

5 4 DIFF

3

3 2 1

2 4

20

6

14

12 10

ALT

4 5

FT 1000

8

2

1

UP

1

DOWN

1

7

CAB ALT

8 9

1000 FT

10

0

0

6

RATE M A N

F A U L T

2

INCR

AUTO

RATE

ALT

CAB SET

BAR

NORM

HOUSING VENT

CAUTION LIGHT

18-PSI BLEED AIR FROM DEICING SYSTEM

DIGITAL COMPUTER

CABIN PRESS ALTITUDE PRESSURE SWITCH

TORQUE

CABIN AIR MOTOR

WEIGHT ON WHEELS POWER LEVERS ADVANCED INPUT CABIN AIR ATMOSPHERE

OUTFLOW VALVE

CABIN A

VENTURI EJECTORINDUCED SUCTION

B

Figure 21-46. Pressurization System Schematic

TECH CHE RR3321B 21

Canada Ltd.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

A schematic of the pressurization control system is shown in Figure 21-46. A schematic of the electrical operation is shown in Figure 21-49.

Automatic When electrical power is applied, the system self-tests, as indicated by illumination of the FAULT light on the selector panel, for approximately two seconds. If the system is faulty, the FAULT light remains illuminated.

feet below ambient. The outflow valve is moved off its fully open stop and starts modulating to be ready to react instantly to the rapid changes, which occur following lift-off. When the landing gear relay is deenergized through the PSEU after lift-off, the valve modulates to maintain pressurization as governed by the computer program.

NOTES

If the self-test is satisfactory, the system is operational and can be controlled in either the CAB SET or NORM mode. In CAB SET, the crew selects the desired cabin cruising altitude, destination altitude (on descent), and cabin altitude rate of change, if desired, using the ALT and RATE knobs on the selector panel. In the NORM mode, only the destination altitude is set, using the ALT knob, and system operation is fully automatic with the information programmed into the computer. In both cases, the BARO knob must be set to correct atmospheric pressure.

On Ground When the airplane is on the ground with the weight on wheels and the power levers are retarded below 80% NH (12° above FLT IDLE), electrical power to pin R of the computer (MSM Chapter 21) is supplied through the normally open contacts of the energized landing gear relay 3261-K1 (the relay is energized through the PSEU). The system is now in the ground mode, and the normal outflow valve is fully open to prevent airplane pressurization.

During Takeoff When the power levers are advanced to 12° above FLT IDLE, power lever switch S1 closes to apply power to pin D of the computer. This, in turn, causes the computer to position the normal outflow valve to pressurize the airplane to 140

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25,000

OPERATIONAL CEILING (25,000 FEET)

20,000

CABIN ALTITUDE—FEET

15,000

UNPRESSURIZED OPERATION

CABIN PRESS CAUTION LIGHT

10,000

MAXIMUM SCHEDULED CABIN ALTITUDE 8,000 FEET NORMAL PRESSURIZATION ENVELOPE 5,000

ULE

SCED

M PR

/NOR

UTO TE A XIMA

RE ESSU

MAXIMUM PRESSURE DIFFERENTIAL SCHEDULE (5.5 PSI)

R

APPO

SL

5,000

10,000

15,000

20,000

25,000

FLIGHT ALTITUDE—FEET

Figure 21-47. Pressurization Envelope

Figure 21-48. Forward Dump Manual Selector

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During Landing

Cabin Pressure Dump

The destination field altitude and atmospheric pressure are set with the ALT and BARO knobs on the selector panel. Pressurization is controlled automatically. If the selected field altitude is set higher than actual field altitude, the airplane lands unpressurized. If the field altitude is set below actual field altitude, the airplane lands pressurized. After landing, cabin altitude returns to field altitude for one minute at the rate selected before dumping cabin pressure to ambient.

Cabin pressure can be dumped by any of the following methods:

A schematic chart of the pressurization envelope is shown in Figure 21-47.

Manual The pressurization level can be controlled with the safety outflow valve by selecting the AUTO–MAN–DUMP switch to MAN and setting the MAN control needle valve to achieve the desired pressurization. Rotating the control clockwise vents the pressure in the safety outflow valve to ambient through the needle valve. This opens the outflow valve, decreasing cabin pressure, which can then be regulated by rotating the MAN control to vary the amount of venting through the needle valve to obtain the desired pressurization level, as shown on the indicator panel.

•

Selecting the AUTO–MAN–DUMP switch to DUMP, causing the normal outflow valve to actuate to the fully open position

•

With CAB SET selected and the system in automatic mode, selecting the ALT knob to a cabin altitude above the flight altitude. This causes the cabin pressure to bleed off at a rate set by the RATE knob.

•

Selecting the forward dump selector to OPEN, causing the safety outflow valve to actuate fully open

•

Switching off bleed air

Forward Dump Manual Selector This selector (Figure 21-48), located on the vertical plane of the copilot’s side console, provides for opening of the safety outflow valve to dump cabin pressure. The selector has positions labeled “NORMAL’’ and “OPEN.’’

Cabin Altitude Warning At a cabin altitude above 10,000 feet, a pressure switch on the avionics rack closes at a cabin pressure of 10 psi. This completes a circuit to illuminate the CABIN PRESS warning light.

If the system is left at AUTO and is operating under automatic control, the resulting decrease in cabin pressure triggers the normal outflow to close when the MAN control knob is rotated to open the safety outflow valve. Thus, as long as the manual selection is for a higher cabin altitude (lower cabin pressure) than set on the automatic control, manual mode overrides the automatic selection.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

P4

P1 CHASSIS GND

C A B

S S

A

NORMAL OUTFLOW VALVE

CABIN SET HI K LO L AUTO J F P M N E G S T U

NORM MAN DUMP

}

S 28V DC L SEC BUS

V SELECTOR PANEL

VALVE-HI (10W) VALVE-LO (10W) +28V DC (25W) CABIN SET-NORMAL +28V DC CABIN DUMP TEST + 28V DC FAULT LAMP RETURN +10V REF POT RETURN BARO CORRECT SIGNAL CABIN RATE SIGNAL LANDING ALT SIGNAL CHASSIS GND PWR GND 28V DC—P.L. AT>12° AFI? 28VDC • WOW CABIN PRESSURE COMPUTER

P39 W e POWER C LEVER SWITCH NO NC S1

D3

J K A F C M N E G S T U V DC B D R

2

D2 D1

(28V DC OUTPUT—WOW) P/O 3261-K1

28V DC

R

3312P2 39

H

CABIN PRESS P/O MASTER CAUTION UNIT

PRESSURE P3 SWITCH C A B ACTUATES AT 10 PSIA (10,000)

A6

RELAY DRIVER

28V DC

LGWOW 1 NGWOW 1 RGWOW 1 LGWOW 2 NGWOW 2 RGWOW 2

LANDING GEAR SENSORS (WEIGHT ON WHEELS)

PROXIMITY SWITCH EELCTRONIC UNIT (PSEU) NOTE: 1. IDENT CODE IS 2131 UNLESS OTHERWISE INDICATED 2 SWITCH (SI) POSITION SHOWN WHEN POWER LEVER IS ADVANCED > 151/2° ABOVE FLT IDLE (SWITCH RELAXED)

Figure 21-49. Pressurization Control—Electrical Schematic

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CHAPTER 21A (SERIES 300A) AIR CONDITIONING CONTENTS Page INTRODUCTION............................................................................................................. 21A-1 BLEED-AIR SYSTEM—SERIES 300 ............................................................................ 21A-3 General ...................................................................................................................... 21A-3 Components Description and Operation ................................................................... 21A-5

System Operation ................................................................................................... 21A-25 BLEED-AIR SYSTEM FLOW CONTROL (MOD 8/1656).......................................... 21A-29 General .................................................................................................................... 21A-29 AIR-CONDITIONING SYSTEM—SERIES 300 .......................................................... 21A-33 General .................................................................................................................... 21A-33 Controls ................................................................................................................... 21A-35 Cooling System ....................................................................................................... 21A-37 Components Description and Operation ................................................................. 21A-39 Operation (Air-Conditioning Package) ................................................................... 21A-48 TEMPERATURE CONTROL SYSTEM........................................................................ 21A-49 General .................................................................................................................... 21A-49 Cabin Temperature Control System ........................................................................ 21A-49 Flight Compartment Temperature Control System ................................................. 21A-49 Components Description and Operation ................................................................. 21A-51 Duct Overheat System ............................................................................................. 21A-55 Temperature Control Operation............................................................................... 21A-57

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CONDITIONED AIR DISTRIBUTION ........................................................................ 21A-61 Cabin Air Distribution ............................................................................................. 21A-61 Flight Compartment Air Distribution (1990 Interior) ............................................. 21A-63 Gasper Air System (1990 Interior) .......................................................................... 21A-65 Air Recirculation System (1990 Interior)................................................................ 21A-67 Avionics Compartment Cooling (1990 Interior) ..................................................... 21A-69 Operation ................................................................................................................. 21A-69

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ILLUSTRATIONS Figure

Title

Page

21A-1

Bleed-Air System Schematic ............................................................................ 21A-2

21A-2

High-Pressure Switches .................................................................................... 21A-4

21A-3

High-Pressure Bleed Shutoff Valve .................................................................. 21A-6

21A-4

Handling Bleed Valve ....................................................................................... 21A-8

21A-5

Precooler Installation...................................................................................... 21A-10

21A-6

Low-Pressure Check Valve ............................................................................ 21A-12

21A-7

Nacelle Shutoff Valve..................................................................................... 21A-14

21A-8

Bleed-Air Overtemperature Switch................................................................ 21A-14

21A-9

Bleed-Air Overpressure Switch ..................................................................... 21A-15

21A-10 Pressure Regulator and Shutoff Valve............................................................ 21A-16 21A-11

Pressure Regulator and Shutoff Valve Schematic .......................................... 21A-18

21A-12 Set-Down Packs Pressure Regulation Schematic........................................... 21A-20 21A-13 Pressure Regulator and Shutoff Valve Schematic (Normal Regulated Flow) 21A-22 21A-14 Bleed-Air System Electrical Schematic—Series 300 .................................... 21A-24 21A-15 Bleed Air System Schematic (Mod 8/1656) .................................................. 21A-28 21A-16 Pressure Regulator and Shutoff Valve Schematic (Mod 8/1656)................... 21A-30 21A-17 Electrical Schematic Diagram, Bleed Simplification (Mod 8/3-1656) .......... 21A-31 21A-18 Air-Conditioning Packs.................................................................................. 21A-32 21A-19 Temperature Control ...................................................................................... 21A-34 21A-20 Air-Conditioning Packs Flow Schematic ....................................................... 21A-36 21A-21 Pack Temperature Control Valves .................................................................. 21A-38 21A-22 Heat Exchanger .............................................................................................. 21A-40 21A-23 Condenser/Mixer............................................................................................ 21A-42

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21A-24 Air Cycle Machine ......................................................................................... 21A-44 21A-25 Compressor Discharge Overtemperature Switch ........................................... 21A-46 21A-26 Cabin Temperature Switch Locations ............................................................ 21A-50 21A-27 Flight Compartment Temperature Switches and Sensors .............................. 21A-50 21A-28 Temperature Controllers................................................................................. 21A-52 21A-29 Flight Attendant Temperature Control ........................................................... 21A-52 21A-30 Caution and Warning Lights .......................................................................... 21A-54 21A-31 Temperature Control Electrical Schematic—Series 300................................ 21A-56 21A-32 Cabin Air Distribution Schematic .................................................................. 21A-60 21A-33 Flight Compartment Air Distribution............................................................. 21A-62 21A-34 Gasper Systems Schematic............................................................................. 21A-64 21A-35 Air Recirculation System and Electrical Schematic ...................................... 21A-66 21A-36 Avionics Compartment Cooling Fan.............................................................. 21A-68

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CHAPTER 21A AIR CONDITIONING (SERIES 300A)

INTRODUCTION This chapter, though titled “Air Conditioning,’’ deals with the environmental systems of the Dash 8, including bleed air, air conditioning, and pressurization. Information is included from Chapter 21, “Air Conditioning’’ and Chapter 36, “Pneumatics,’’ of the 300 Series Maintenance Manual. The material in this chapter is oriented toward the line mechanic. All values expressed throughout this chapter, such as for pressure, temperature, flow rates, and time, are used only for their illustrative meanings. Actual values may differ and must be obtained from the pertinent sections of the Maintenance Manual.

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1A

1B

6

3 2 HP 2 NO. 2 ENGINE

5 ECU

LP

TO DEICING SYSTEM

4 7 AIR CONDITIONING

11 FROM CABIN SUPPLY DUCT

OFF

RECIRC CABIN

1

BLEED

12

RECIRC CABIN F/C

2

8

NORM 40 20 °C 0 DUCT TEMP

13

60

9

80 100

MIN BLEED

MAX

#2 BLEED HOT

CABIN CAB DUCT

10 FC DUCT

14

GAUGE

15

OFF MAN AUTO COOL

WARM

F/A CABIN

PACKS

TEMP CONTROL

COOL

WARM

TO NO. 1 AIRCONDITIONING PACK

FLT COMP

15 10

FROM NO. 1 ENGINE BLEED AIR SYSTEM (SIMILAR TO NO. 2 SYSTEM)

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

HIGH-PRESSURE SWITCHES CHOKING VENTURI HIGH-PRESSURE SHUTOFF VALVE LOW-PRESSURE CHECK VALVE HANDLING BLEED VALVE (HBOV) PRECOOLER NACELLE SHUTOFF VALVE BLEED OVERPRESSURE SWITCH OVERTEMPERATURE SWITCH WING ISOLATION CHECK VALVE SERVO AIR FILTER FLOW CONTROL SERVO SET-DOWN LIMITER SHUTTLE VALVE PRESSURE REGULATOR AND SHUTOFF VALVE

TO NO. 2 AIRCONDITIONING PACK

Figure 21A-1. Bleed-Air System Schematic

21A-2

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BLEED-AIR SYSTEM— SERIES 300

NOTES

GENERAL The bleed-air system in each nacelle consists of an optional precooler with a removable bypass section, two high-pressure switches, a high-pressure shutoff valve, high-pressure choking venturi, low-pressure check valve, nacelle shutoff valve, and duct overpressure switch (Figure 21A-1). The high- and low-pressure engine bleeds join downstream. A bleed duct overtemperature switch and a wing isolation check valve route bleed air to two air-conditioning packs in the aft fuselage. Two pressure regulator and shutoff valves are controlled by a single flow control servo unit, set-down limiter, and shuttle valve. Each nacelle bleed-air system is controlled by a two-position bleed switch on the AIR CONDITIONING panel (Figure 21A-1). The BLEED flow control knob adjusts both pressure regulators, controlling the flow rate. The engines have a high flow and pressure ratio. To achieve the higher first-stage pressure at rated power settings, it is necessary to spill excess low-pressure air with the engine at idle. This is achieved with a bleed-off valve (HBOV) at the low-pressure bleed port. The valve, installed by the engine manufacturer, is actuated by a torque motor controlled by the engine electronic control (EEC). It is open during ground idle and ramped closed as engine rpm increases. The valve also opens at flight idle near sea level, but not at operating altitude (25,000 feet) except during engine restart.

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NOTE: SWITCHES IN BOTH LOCATIONS ARE IDENTICAL EXCEPT FOR PRESSURE SETTINGS.

Figure 21A-2. High-Pressure Switches

21A-4

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COMPONENTS DESCRIPTION AND OPERATION

NOTES

High-Pressure Switches Two high-pressure switches (Figure 21A-2), located on the wing front spar outboard of the nacelle, are connected to a tapping in the outlet duct from the engine high-pressure bleed port. Each switch is a normally closed pneumatic pressure-sensing switch containing diaphragm-operated electrical contacts which are connected into the circuit to the high-pressure shutoff valve when the BLEED switch is selected to BLEED. The No. 1 switch is set at 65 psig; the No. 2 switch is set at 55 psig.

Operation In ground taxi mode with weight on wheels (WOW), the No. 1 switch (65 psig) is activated, allowing the airplane bleed air to remain on high pressure through taxiing. A sensed pressure of 65 psig increasing moves the diaphragm to open a set of contacts in the switch; the contacts close when the sensed pressure drops to 65 psig. This switch is also activated when the deice AIRFRAME AUTO switch is selected to SLOW or FAST. In the flight mode (no WOW signal), the only difference from the ground mode is that the No. 2 (55 psig) switch is activated instead of the No. 1 switch. However, when the deice AIRFRAME AUTO selector is positioned to SLOW or FAST, bleed air will be controlled by the No. 1 switch.

High-Pressure Bleed-Air Venturi A choking venturi (Figure 21A-1) installed in the HP bleed port restricts bleed-air flow to a maximum of 10%. This restriction prevents engine bleed from reaching damaging proportions, such as both air-conditioning packs operating from a single engine or in the event of a duct rupture.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SOLENOID ACTUATOR HOUSING

ELECTRICAL CONNECTOR ACTUATOR COVER

ACCESS TO FILTER

FLOW CONTROL VALVE

SOLENOID SHOWN ENERGIZED FROM PRESSURE SWITCH

A B C

SOLENOID SHOWN DEENERGIZED FROM PRESSURE SWITCH

A B C

CONNECTOR

CONNECTOR VALVE OPEN

FILTER

VALVE OPEN

FILTER DIRECTION OF FLOW

DIRECTION OF FLOW

TWO-INCH DISC

Figure 21A-3. High-Pressure Bleed Shutoff Valve

21A-6

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High-Pressure Bleed Shutoff Valve

NOTES

The high-pressure (HP) bleed shutoff valve (Figure 21A-3) is line-mounted in the hot air duct from the high-pressure bleed port on the engine. It is a normally closed, solenoid-controlled pneumatically operated valve used to control the flow of HP bleed air from the engine. The valve consists of a solenoid and a pneumatic actuator which is mechanically linked to a butterfly valve. When the solenoid is energized, servo pressure from upstream is applied to the pneumatic actuator to open the butterfly. To ensure correct installation of the valve in the system, a flow direction arrow is provided on the valve body.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DUCT TUBE

COVER

RESTRICTOR P2.4

P2.4

MANIFOLD

TO EEC

SERVO CONTROLLED P2.4 BLEED

PISTON SERVO VALVE

ELECTRICAL HARNESS CONNECTION

P2.5

OPEN POSITION

P2.5

CLOSED POSITION

Figure 21A-4. Handling Bleed Valve

21A-8

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Handling Bleed Valve Control

NOTES

The handling bleed valves (HBV) (Figure 21A-4), located on each engine, are connected with and control the outflow of the associated engine P 2.5 bleed air overboard through a precooler in the bleed-air system The HBV is used to prevent compressor surge and stall and to reduce noise level during steady operation of the engine. It consists of a housing, a piston in a ported sleeve, a cover, a servo valve, and a manifold with a restrictor. The HBV piston is controlled by P 2.5 air on its underside and by P 2.4 air on the opposite side. P 2.4 air is ducted from a diffuser pipe through the restrictor and manifold, then into the servo valve. On a signal from the EEC, the flapper nozzle in the servo valve opens and vents P 2.4 air. P 2.4 air in the HBV undergoes a pressure drop, which allows P 2.5 air to seat the piston. The valve is then open, and P 2.5 air vented into an airframe tube is ducted to the airframe precooler (to reduce P 3 bleed temperature) and is then vented overboard. On a signal from the EEC, the flapper nozzle in the servo valve closes; P 2.4 air stops venting, and its pressure stabilizes to seat the piston against P 2.5 air. The piston stops P 2.5 air from entering the valve and venting. The valves are controlled by the corresponding engine ECU through a potentiometer on the power lever layshaft.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FWD

PRECOOLER UNIVERSAL JOINT

HP SHUTOFF VALVE

TAP OFF TO HP SWITCHES

HP VENTURI

Figure 21A-5. Precooler Installation

21A-10

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Precooler

NOTES

The precooler (Figure 21A-5) is a steel heat exchanger mounted above the engine in the high-pressure (HP) bleed-air supply line. Its purpose is to cool the hot HP engine bleed air to an acceptable temperature for the aluminum heat exchangers of the two rear fuselage air-conditioning packs. The cooling effect is achieved by the handling bleed-off (HBOV) air flowing across the precooler and out the zone 2 exhaust louver on the top of the engine nacelle.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

HOUSING ASSEMBLY

SEAT

OPEN POSITON

PISTON

CLOSED POSITON

HANDLING BLEED VALVE

INTERCOMPRESSOR CASE (REF) P2.5

Figure 21A-6 Low-Pressure Check Valve

21A-12

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Low-Pressure Check Valve

NOTES

The low-pressure check valve (Figure 21A-6) is mounted in the low-pressure (LP) bleed port. Its prime function is to isolate LP bleed air from the bleed-air system when the HP bleed system is operating. The LP bleed port does not have a venturi like the HP system, but the check valve has a restrictor built in to perform the same function. The valve limits LP bleed to a maximum of 10% of bleed flow.

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Figure 21A-7. Nacelle Shutoff Valve

Figure 21A-8. Bleed-Air Overtemperature Switch

21A-14

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Nacelle Shutoff Valve

Bleed-Air Overpressure Switch

The nacelle shutoff valve is installed in the outlet duct of the nacelle bleed-air system (Figure 21A-7). The function of the valve is to shut off bleed air to the air-conditioning packs when they are not selected and when the engine high-pressure bleed port overpressure or overtemperature limits are exceeded.

The wing bleed overpressure switch is installed in the bleed-air delivery duct, downstream from the nacelle shutoff valve (Figure 21A-9). If bleed-air pressure exceeds 75 psig, the switch completes a circuit to close the nacelle shutoff valve, isolating the affected wing and illuminating the appropriate BLEED HOT caution light.

Overtemperature Switch

NOTES

The overtemperature switch (Figure 21A-8) is installed in the bleed-air delivery duct downstream of the nacelle shutoff valve. It consists of a normally open thermal switch with a bimetallic element and closes when air temperature in the duct exceeds 550 +10° F (287 +5° C) and illuminates the appropriate BLEED HOT caution light.

TECH CHECK

Figure 21A-9. Bleed-Air Overpressure Switch

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONTROL COVER SOLENOID

ACCESS TO FILTER

ACTUATOR AND TUBE

CONTROL HOUSING

PRESSURE TEST PORT

ACTUATOR HOUSING

FLOW CONTROL VALVE

NO. 2 VALVE

NO. 1 VALVE

Figure 21A-10. Pressure Regulator and Shutoff Valve

21A-16

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Pressure Regulator and Shutoff Valve

NOTES

There are two pressure regulator and shutoff valves (PRSV) installed in the bleed-air ducting in the rear fuselage aft of the rear pressure bulkhead (Figure 21A-10). One is mounted in the bleed-air duct to the No. 1 (rear) air-conditioning pack while the other is mounted in the bleed-air duct to the No. 2 (forward) airconditioning pack. The purpose of the valves is to shut off bleed air to the air-conditioning packs when air conditioning is not required and to regulate the flow of bleed air to the respective air-conditioning systems.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FROM FLOW CONTROL SERVO VIA SET-DOWN LIMITER PRESSURE -SENSING CONTROL SIGNAL CAPSULE

ANEROID

STABILIZER UNIT RESTRICTOR 28 VDC

AMB SUMMING BAR

REGULATING CAPSULE

SOLENOID (SHOWN DEENERGIZED)

PNEUMATIC ACTUATOR

TEST PORT OPEN

FILTER

AMB

NOTE: BUTTERFLY VALVE OPEN

NO. 2 AIR-CONDITIONING SYSTEM PRSOV SHOWN, NO. 1 PRSOV SIMILAR

PRESSURE-SENSING CONTROL ANEROID SIGNAL CAPSULE

STABILIZER UNIT RESTRICTOR 28 VDC

AMB SUMMING BAR

REGULATING CAPSULE

SOLENOID (SHOWN ENERGIZED)

PNEUMATIC ACTUATOR

TEST PORT OPEN

FILTER

AMB

Figure 21A-11. Pressure Regulator and Shutoff Valve Schematic

21A-18

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Pneumatic Actuator Operation

NOTES

Pressure for valve operation is from upstream of the valve and is routed to both sides of the pneumatic actuator piston (Figure 21A-11). The close side is spring-assisted; the open side has a larger area to overcome the combined force of the spring and air pressure. The amount of pressure to the open side is controlled by the pressure-sensing control, which consists of four pneumatic inputs to a summing bar that controls pressure venting to establish open side pressure. These inputs are: • An actuator sensing downstream pressure • A stabilizer unit controlled by a restricted line • An aneroid that senses pressure altitude pressure changes • A servoactuator that receives the output of the flow control servo

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

REGULATED PRESSURE FROM PRSOV

LAMINAR ORIFICE

ADJUSTMENT SCREW SET AT 6 PSI

CONTROL LEVER

POPPET

CABIN PRESSURE FROM CABIN AIR SUPPLY DUCT SET-DOWN (RESET) PRESSURE FROM FLOW CONTROL SERVO

SET-DOWN LIMITER SCHEMATIC

18 PSI DEICE PRESS

FLOW CONTROL SERVO

FILTER

DORSAL BLEED PIPE SET-DOWN SIGNAL (FROM APU IF FITTED)

SET-DOWN LIMITER PRSOV CABIN SUPPLY SIGNAL

PRSOV REG. PRESS SIGNAL

SHUTTLE VALVE AFT NO. 1 PACK CABIN FWD NO.2 PACK FLT COMPT

Figure 21A-12. Set-Down Packs Pressure Regulation Schematic

21A-20

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Flow Control Servo

NOTES

The flow control servo (Figure 21A-12), located in the aft fuselage aft of the rear pressure dome, meters deice system air to the signal port in the PRSVs via a set-down limiter and shuttle valve. A DC signal from the bleed flow control positions a torque motor to reduce the signal pressure.

Shuttle Valve The shuttle valve is a single ball-type valve that receives output pressure from PRSVs. The valve with the highest output pressure influences the set-down limiter, modifying venting of the limiter.

Set-Down Limiter The set-down limiter is connected into the signal line from the flow control servo to the two PRSVs via a tee fitting. The purpose of the set-down limiter is to limit the minimum output pressure of the two PRSOVs. If the higher of the two valves’ downstream pressure is not more than 6 psi above cabin pressure, the set-down limiter vents signal pressure to maintain sufficient flow to pressurize the cabin. The set-down limiter allows the NORM (center) position of the bleed flow control knob to be used, thus providing two-thirds of the total output of the two air-conditioning packs. This setting is used to economize on bleed flow without requiring continuous adjustment.

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BLEED FLOW CONTROL

ELECTROPNEUMATIC SERVO TORQUE MOTOR

BLEED MAX

MIN

FILTER 28 VDC L AND R SEC BUS

18 PSI FROM DEICE SYSTEM PRESSURE SENSING CONTROL

SET DOWN LIMITER

SERVOACTUATOR

STABILIZER UNIT

ANEROID RESTRICTOR 28 VDC SOLENOID (SHOWN DEENERGIZED)

AMB SUMMING BAR

PNEUMATIC ACTUATOR

SENSING ACTUAOR TEST PORT CLOSE

FILTER AMB

LEGEND BLEED AIR REGULATED AIR REGULATED AIR (18 PSI) AMBIENT

BUTTERFLY VALVE

Figure 21A-13. Pressure Regulator and Shutoff Valve Schematic (Normal Regulated Flow)

21A-22

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PRSV Operation The PRSV (Figure 21A-13) is fail-safe open in the event of an electrical failure. Selecting the PACK switch on the AIR CONDITIONING panel to OFF energizes the solenoid to close off the supply of air pressure to the open side of the actuator (Figure 21A13) . The spring and air pressure to the close side moves the actuator, closing the butterfly. Selecting the PACK switch to AUTO or MANUAL deenergizes the solenoid, allowing air pressure to open the butterfly valve.

e n s m o v e m e n t o f t h e s e n s i n g a c t u a t o r. Pneumatic actuator and butterfly valve operation is the same as previously described. To reduce the amount of air being bled from the engine as altitude increases, an aneroid expands to vent more air pressure. The aneroid contracts as altitude decreases, venting less air pressure.

NOTES

With the PACK switch in AUTO or MANUAL, the valve regulates bleed-air flow to the air-conditioning system. This is accomplished by the pressure-sensing control venting a measured quantity of pneumatic actuator openside air pressure overboard; the amount is dependent on three interrelated factors: • Bleed flow rate required • Altitude • Bleed-air pressure sensed downstream of the valve If a high flow rate is selected by the BLEED control knob, the torque motor is positioned to vent a minimum amount of pressure from the open side of the pneumatic actuator. The remaining pressure overcomes the spring to drive the butterfly toward the open position. If a low flow rate is desired, the torque motor is positioned to vent a small amount of servo pressure overboard, the remaining pressure acting on the servoactuator to vent more pressure from the open side of the actuator. Spring force overcoming the servo pressure drives the butterfly toward the closed position. With the flow rate set, air venting is controlled by pressure sensed downstream of the valve. Pressure increase acts on a sensing actuator to vent more air; decreasing pressure results in the sensing actuator returning to normal and less air being vented. The stabilizer unit damp-

Revision 2

21A-23

21A-24 NO. 1 HIGH PRESSURE SWITCH NO

NC

HIGH PRESSURE BLEED SHUT-OFF VALVE

NO. 2 HIGH PRESSURE SWITCH NO

NACELLE SHUT-OFF VALVE (OPEN)

OVERPRESSURE SWITCH NC

NC

NO CONTACT CLOSES P >75 PSI

B A

P19

A B C

P7

A

B

B

P3

A

P13

B

C

CLOSES AT T > 550 ±10°

P13

A

B

P5

OPENS WHEN PRESSURE > 55 PSI

A B CR9 C

NC

A3

D

A2 A1

SEE SHEET 2

B3 B2

NC A3

B1 D3 X1

D2

X2

D1

3

{

X1 X2 6121-K16

HBOV 0VERIDE BLEED SYSTEM 1 S3

1 2 3

OFF

4 5 6

BLEED ON BLEED SYSTEM CONT 1 5A (B5) 28V DC LEFT MAIN BUS CB1

CR 11

C1

K1 HIGH PRESSURE BLEED CONTROL RELAY (NORMALLY CLOSED)

2

4

A1 B3 CR6

B2

CR1

C3

C2

2

CR5

A2

NO

B1 NO 5 SEC TIME DELAY ON OPERATE B3 (AIR)

CR2 X1 X2 K3 TIME DELAY ON RELEASE A3 A2 A1 C1 X1 X1

B2

CR 19

B1 (GND)

X1 X2

A3 A2 A1

D1

X1

R3 X2

D3 3010-K2

E

K6

5 AUXILIARY LANDING RELAY

3261-K4 AIRFRAME AUTOMATIC SWITCH 1 (OFF) (SLOW) 2

2C1

CR20 (FAST) 3

3010-S3 4

AIRFRAME DEICE 28V DC (E8) 7.5 R. SEC CB4 BUS

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NO

OPEN WHEN PRESSURE > 65PSI

OVERTEMPERATURE SWITCH

Canada Ltd.

Revision 2

Figure 21A-14. Bleed-Air System Electrical Schematic—Series 300 (Sheet 1 of 2)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SYSTEM OPERATION

NOTES

Selecting the BLEED switch to BLEED (Figure 21A-14) deenergizes the nacelle shutoff valve and both pressure regulator valves open and provides 28 VDC to the solenoid of time delay relay K3 through the normally closed contacts of the high-pressure bleed control relay (MSM Chapter 21). Power is also applied to time delay relay 3010K2 through the normally closed contacts of time delay relay K3 to control terminal C1 of deice delay relay 3010-K2. With both high-pressure switches closed, deice time delay relay 3010-K2 energizes and its normally closed contacts open to disconnect the ground from the deice system low-pressure warning switch. This prevents the DE-ICE PRESS caution light from illuminating during high-pressure shutoff valve operation. Five seconds from BLEED selection, time delay relay K3 energizes, and the high-pressure shutoff valves are energized open. With the valves open and the engines operating, high-pressure bleed air flows to the air-conditioning packs through the choking venturi, nacelle shutoff valves, and pressure regulator and shutoff valves. Five seconds from the opening of the normally closed contacts of time delay relay K3, power is disconnected from deice delay relay 3010-K2 terminal C1. The relay is deenergized, reconnecting the ground to the deice system low-pressure warning switch.

Revision 2

21A-25

21A-26 NO. 1 HIGH PRESSURE SWITCH NO

NC

C B A

NO

P18

NACELLE SHUT-OFF VALVE

OVERPRESSURE SWITCH

NC

A B C

OVERTEMPERATURE SWITCH

NO CONTACT CLOSES AT P >75 PSI A

P8

B

B

P9

A

P4

C B A

P12

CLOSES AT T > 550 ±10° A

B

P6

OPENS WHEN PRESSURE > 55 PSI

'B' 'A' 'E' 'D' SEE SHEET 1

A3 A2 3312-P3 'C'

12

A1 LOGIC CIRCUIT

NO.1 BLEED HOT

B2

B3 B1

D3 X1

D2

C2

C3 X1 X2

3

6121-K15

3312-P31 7

LOGIC CIRCUIT

NO.2 BLEED HOT

MASTER CAUTION PANEL

28V DC CAUTION LIGHTS POWER

HBDV OVERRIDE OFF

BLEED SYSTEM CONT 2 (R5) 5A 28V DC RIGHT MAIN BUS CB2

CR12

2

X2

D1 C3

K2 HIGH PRESSURE BLEED CONTROL RELAY

1 2 3 4 5 6

BLEED ON 2

BLEED SYSTEM 2 S4

CR10

RIGHT BLEED NOTES: 1. IDENT CODE IS 2121, UNLESS OTHERWISE SPECIFIED. 2 FOR CONTINUATION, SEE CHAPTER 73, ENGINE ELECTRONIC CONTROL SYSTEM—ENGINE 1 AND 2. 3 FOR CONTINUATION, SEE CHAPTER 61, PROPELLER AUTO FEATHER CONTROL SYSTEM. 4 FOR CONTINUATION, SEE CHAPTER 30, AIRFRAME DE-ICE AND INDICATION. 5 FOR CONTINUATION, SEE CHAPTER 32, PROXIMITY SWITCH SYSTEM.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

OPEN WHEN PRESSURE > 65PSI

HIGH PRESSURE BLEED SHUT-OFF VALVE

NO. 2 HIGH PRESSURE SWITCH

Canada Ltd.

Revision 2

Figure 21A-14. Bleed-Air System Electrical Schematic—Series 300 (Sheet 2 of 2)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

When either nacelle upstream high-pressure bleed air reaches 65 psi on the ground or 55 psi in the air, the associated high-pressure switch opens and disconnects the ground from time delay relay K3 and deice time delay relay 3010-K2. Relay K3 deenergizes both highpressure shutoff valves closed and reconnects power to relay 3010-K2 terminal C1, energizing the relay and disconnecting the ground from the deice low-pressure warning switch. Low-pressure bleed air is then supplied to the system.

NOTES

Airborne with the AIRFRAME AUTO switch in SLOW or FAST, the bleed-air system remains in the high-pressure mode, controlled by the No. 1 high-pressure switch (65 psig). If an overpressure condition occurs, the wing bleed overpressure switch closes at 75 psi to energize the nacelle shutoff valve closed, isolating the affected bleed-air system.

Revision 2

21A-27

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

1

1

6

3 2 HP 2 NO. 2 ENGINE

5

LP

TO DEICING SYSTEM

4 7 AIR CONDITIONING OFF

RECIRC CABIN

13 14

OFF 1

BLEED

0 DUCT TEMP

14

RECIRC F/C

2

8

NORM 40 20 °C

13

60

9

80 100

MIN BLEED

MAX

CABIN CAB DUCT

#2 BLEED HOT

12

12

10 FC DUCT

GUAGE

COOL

WARM

F/A CABIN

OFF AUTO MAN PACKS

TEMP CONTROL

11 COOL

15

16

WARM

FLT COMP

15 10 16

1. HIGH-PRESSURE SWITCHES 2. CHOKING VENTURI 3. HIGH-PRESSURE SHUTOFF VALVE FROM NO. 1 ENGINE 4. LOW-PRESSURE CHECK VALVE BLEED AIR SYSTEM 5. HANDLING BLEED VALVE (HBOV) (SIMILAR TO NO. 2 SYSTEM) 6. PRECOOLER 7. NACELLE SHUTOFF VALVE 8. BLEED OVERPRESSURE SWITCH 9. OVERTEMPERATURE SWITCH 10. WING ISOLATION CHECK VALVE 11. BLEED FLOW CONTROL 12. SOLENOID VALVE 13. LEE JET RESTRICTOR 14. SIGNAL PRESSURE PORT 15. PRESSURE REGULATOR AND SHUTOFF VALVE 16. REGULATED PRESSURE PORT

TO NO. 1 AIRCONDITIONING PACK

TO NO. 2 AIRCONDITIONING PACK

Figure 21A-15. Bleed-Air System Schematic (Mod 8/1656)

21A-28

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

BLEED-AIR SYSTEM FLOW CONTROL (MOD 8/1656)

NOTES

GENERAL The modification introduces a simplified method of controlling bleed flow to the air conditioning (A/C packs). Two solenoid valves (one each for the PRSOV of the A/C packs), together with a Lee restrictor, are installed in a line joining each PRSOV regulated pressure port to its signal pressure port. The solenoid valves are a normally closed (deenergized) and, when in the closed position, no air flow reaches the PRSOV signal port and it senses the ambient pressure and enforces a maximum bleed flow to the A/C pack. When the solenoid valves are energized open, air flow is allowed through the restrictors to the PRSOV signal pressure port and it senses higher than ambient pressure and enforces a decreased bleed flow to the A/C pack.

Revision 2

21A-29

21A-30

13

14

PRESSURE SENSING CONTROL

STABILIZER UNIT 15 ANEROID RESTRICTOR 28V DC

12 11

SOLENOID (SHOWN DEENERGIZED)

AMB SUMMING BAR

SENSING ACTUATOR

PNEUMATIC ACTUATOR

TEST PORT CLOSE

FILTER AMB

16 NORM MIN

MAX

BLEED BUTTERFLY VALVE

Canada Ltd.

Revision 2

Figure 21A-16. Pressure Regulator and Shutoff Valve Schematic (Mod 8/1656)

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SERVO ACTUATOR

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

The system is controlled by a new flow control selector switch on the overhead air conditioning panel in the flight compartment. The switch has three separate positions (MIN–NORM–MAX) and the effect of each position is as follows:

NOTES

MIN—Both solenoid valves are energized (open), the No. 2 (fwd) A/C pack PRSOV energized (closed) and the No. 1 (rear) A/C pack PRSOV deenergized (open), resulting in the rear A/C pack receiving decreased bleed-air flow. NORM—Both solenoid valves are energized (open), both No. 1 and No. 2 A/C pack’s PRSOV are deenergized open, resulting in both A/C Packs receiving decreased bleed-air flow. MAX—Both solenoid valves are deenergized (closed) both No. 1 and Nol. 2 A/C pack’s PRSOV are deenergized (open), resulting in both A/C packs receiving maximum bleed-air flow.

FWD PACK PR AND SOV

FWD PACK/AFT PACK NORMAL FLOW SOLS.

N.O

N.C

N.C

R2 R1

NORMAL K2

K1 5 SEC

MIN

MAX

5 SEC DELAY

S1 DELAY BLEED FLOW CONTROL

Figure 21A-17. Electrical Schematic Diagram, Bleed Simplification (Mod 8/3-1656)

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NO. 1 PACK

NO. 2 PACK

Figure 21A-18. Air-Conditioning Packs

21A-32

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR-CONDITIONING SYSTEM—SERIES 300

NOTES

GENERAL Conditioned air from two air-conditioning packs is ducted into the fuselage at the center of the rear pressure bulkhead (Figure 21A18). The No. 1 (rear) pack supplies conditioned air to the cabin; the No. 2 (forward) pack supplies conditioned air to the flight compartment and cabin. The flight compartment duct is routed under the baggage compartment and cabin, where it splits into two ducts to feed the pilot’s and copilot’s subsystems. The cabin supply duct is routed under the baggage compartment, where it splits into an upper and lower duct for each side of the fuselage. The upper duct supplies air to the cabin dado panel grills; the lower duct supplies, via sidewall risers, the upper cabin air outlets, as well as the cabin, lavatory, and flight attendant’s station gaspers. The flight compartment duct supplies air to sidewall grills and gaspers in the flight compartment. Flight compartment air is also used for side window demisting. On airplanes with SOO 8069, conditioned air can be supplied to the cabin and flight compartment through an eight-inch universal ground air service connector located on the left or right side of the rear fuselage. The air supply from the ground connection is routed through a short flexible duct to join with the recirculated air duct in the rear fuselage. The air is then routed via the condenser/mixer to the cabin and flight compartment. A butterfly-type check valve is lined-mounted in the ground conditioned air duct to prevent recirculated air or air-conditioning pack air from spilling overboard through the ground air connector.

Revision 2

21A-33

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RECIRCULATION FAN SWITCH RECIRC — ACTIVATES THE RESPECTIVE RECIRCULATION FAN ON. DUCT TEMPERATURE INDICATOR INDICATES AIR TEMPERATURE IN THE CABIN AND SUPPLY DUCTS FROM BOTH AIRCONDITIONING PACKS.

AIR CONDITIONING HBDV OFF

DUCT TEMPERATURE SELECTOR (ROTARY SELECTION) CAB DUCT— INDICATES AIR TEMPPERATURE IN SUPPLY DUCT FROM AFT PACK.

RECIRC CABIN

OFF

1

0 DUCT TEMP

60 80

100

MIN BLEED

AUTO — TURNS ON ASSOCIATED PACK FOR AUTOMATIC TEMPERATURE CONTROL. MAN — TURNS ON ASSOCIATED PACK; HOWEVER, TEMPERATURES ARE SET WITH MANUAL TEMPERATURE CONTROL SWITCHES.

MAX

CABIN CAB DUCT

FC DUCT— INDICATES TEMPERATURE IN SUPPLY DUCT FROM FORWARD PACK.

OFF— SELECTS ASSOCIATED PACK OFF.

RECIRC F/C

2

NORM 40 20 °C

CABIN— INDICATES CABIN AIR TEMPERATURE.

PACK CONTROL SWITCHES LEFT SWITCH CONTROLS AFT PACK (CABIN); RIGHT SWITCH CONTROLS FORWARD PACK (FLIGHT COMPARTMENT).

BLEED

FC DUCT

MANUAL TEMPERATURE CONTROL SWITCH • PROVIDES MANUAL CONTROL WHEN PACK SWITCH IS SELECTED TO MAN • HOLDING SWITCH TOWARD COOL OR WARM DIRECTS PACK BYPASS VALVES IN SELECTED DIRECTION.

GUAGE

COOL

WARM

F/A CABIN

OFF AUTO MAN PACKS

TEMP CONTROL

COOL

WARM

FLT COMP

AUTOMATIC TEMPERATURE CONTROL (ROTARY ACTION) • PROVIDES AUTOMATIC TEMPERATURE CONTROL FOR ASSOCIATED COMPARTMENT WHEN PACK SWITCH IS SELECTED TO AUTO. • CABIN TEMPERATURE CONTROL CAN BE DELEGATED TO THE FLIGHT ATTENDANT BY TURNING COUNTERCLOCKWISE TO F/A POSITION (DETENT).

Figure 21A-19. Temperature Control

21A-34

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONTROLS

NOTES

Controls on the AIR CONDITIONING panel (Figure 21A-19), located on the overhead console in the flight compartment, include a TEMP gage to monitor the temperature in the cabin compartment, cabin or flight compartment ducts, and switches for cabin and flight compartment temperature control. Temperature in the cabin and flight compartment can be controlled either auto matically or manually by selection of the two PACKS switches marked OFF—AUTO—MAN. The two manual temperature controls are momen tarily-on toggle switches sharing the same COOL-WARM legend with the automatic temperature control rot a r y va r i a b l e r e s i s t o r s . T h e m a n u a l a n d automatic temperature control switches are wired into the appropriate temperature controller. The cabin temperature can also be controlled through a flight attendant control rotary switch located at the forward flight attendant’s station. Power for automatic control is supplied through the temperature controller.

Revision 2

21A-35

21A-36

RAM AIR

BAFFLE

ENGINE BLEED AIR CHECK VALVE

DUCT TEMP SENSOR

ZONE/DUCT OVERTEMP DUCT TEMP SENSOR SENSOR

TAILCONE PRESSURE RELIEF

RAM AIR OVERBOARD PACK TEMP CONTROL VALVE

COMPRESSOR DISCHARGE OVERTEMP SWITCH

HEAT EXCHANGER

CONDENSER

MIXING BOX FILTER

DUCT DUCT OVERTEMP TEMP SENSOR SENSOR

CHECK VALVE

HEAT EXCHANGER

CONDENSER

MIXING BOX FILTER WATER NOZZLE

WATER NOZZLE SILENCER RAM AIR AIR CYCLE MACHINE (ACM)

NO. 1 RECIRCULATION FAN

SILENCER AIR CYCLE MACHINE (ACM)

NO. 2 AIR-CONDITIONING PACK OVERBOARD DRAIN

RAM AIR

NO. 1 AIR-CONDITIONING PACK OVERBOARD DRAIN

Figure 21A-20. Air-Conditioning Packs Flow Schematic

Canada Ltd.

Revision 4—July 1995

NO. 2 RECIRCULATION FAN

PACK TEMP CONTROL VALVE

COMPRESSOR DISCHARGE OVERTEMP SWITCH

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

APU BLEED AIR BLEED PRESSURE REGULATOR AND SHUTOFF VALVE

BLEED PRESSURE REGULATOR AND SHUTOFF VALVE

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COOLING SYSTEM

NOTES

The cooling system consists of two air-conditioning packs mounted in tandem in the rear fuselage aft of the rear pressure dome (Figure 21A-20). Both packs are supplied with bleed air from both engines and cools it before distribution to the systems. When heat ing is required, the bleed air is allowed to bypass the refrigeration section of the se lected sys tem’s pack. If cooling is required, the selected system’s temperature control valves close, routing the bleed air through its refrigeration section. Temperature of the air supplied by either pack is controlled automatically or manually from the air-conditioning panel in the flight compartment. In the event of a single pack failure, the remaining pack supplies conditioned air for all systems. If both packs fail, ram air is supplied to the cabin and flight compartment for ventilation. Each air-conditioning pack consists of an air cycle machine (ACM), a heat exchanger, a condenser/mixer, electrically operated pack temperature control valves, and a compressor discharge overtemperature switch.

Revision 2

21A-37

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

THROTTLE BUTTERFLY VALVE

PACK BYPASS BUTTERFLY VALVE

ACTUATOR

MOUNTING PLATE

COLD FULL-TRAVEL MICROSWITCH 2.5-INCH THROTTLE VALVE LINKAGE OPEN VALVE POSITION

1.5-INCH BYPASS VALVE

CLOSED CCW STOP

CLOSED ACTUATOR STROKE

CW STOP

TE RR33

FUNCTIONAL SCHEMATIC

Figure 21A-21. Pack Temperature Control Valves

21A-38

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION AND OPERATION

NOTES

Pack Temperature Control Valve Each of the two air-conditioning packs has a pack temperature control valve (Figure 21A21) which consists of a 1.5-inch-diameter pack bypass butter fly valve operated in opposition to a 2.5-inch diameter throttle butterfly valve by an electric actuator. The two valves are mechanically linked to the actuator, and all three components are installed on a common mounting plate attached to the heat exchanger of each air-conditioning pack. Each pack bypass valve is located in a bypass duct from the bleed-air duct, and the throttle valve is located in the outlet duct from both primary heat exchangers. The control valves regulate the temperature of the air discharged from the air-conditioning packs by opening and closing the appropriate pack bypass and throttle valves. The valves are controlled by electrically operated actuators which modulate in response to automatic signals from the cabin temperature controller, or from electrical inputs from manually operated switches. The mechanical linkage connecting the two valves to their actuator is designed to effect a sequence of valve opening and closing in response to rotation of the actuators. This provides a complete temperature range from maximum cooling to maximum heating. For maximum cooling, the appropriate pack bypass valve is closed and the respective throttle valve is open, allowing the total flow from the engine bleed-air system to pass through a dual heat exchanger and air cycle machine. When less cooling is required, the actuator moves the throttle valve toward the closed position and the bypass valve toward open to allow a proportional amount of the airflow to bypass the refrigeration circuit. At maximum heat, the throttle valve is fully closed, and the pack bypass valve is fully open.

Revision 2

21A-39

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RAM-AIR OUTLET HEADER

VENT DOORS

BLEED BYPASS TO MIXING BOX INLET DUCT (FROM ENGINE BLEED OR APU [SOO 8062])

RAM-AIR OUTLET

PACK TEMPERATURE CONTROL VALVE MOUNTING BRACKET PRIMARY

SECONDARY

OUTLET DUCT (TO ACM COMPRESSOR)

RAM-AIR INLET INLET DUCT (FROM ACM COMPRESSOR) OUTLET DUCT (TO ACM TURBINE)

TRANSITION DUCT RAM-AIR INLET HEADER

WATER NOZZLE

Figure 21A-22. Heat Exchanger

21A-40

Revision 4—July 1995

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Heat Exchanger

NOTES

Two identical heat exchangers are located in the fuselage aft of the rear pressure dome (Figure 21A-22). Each consists of a brazed aluminum core which is divided into separate primary and secondary portions. The primary heat exchanger inlet duct is connected to the engine bleed-air supply; the outlet duct is connected to the air cycle machine compressor inlet. The secondary heat exchanger inlet duct is connected to the air cycle machine compressor outlet; the outlet duct is connected to the air cycle machine turbine inlet via the condenser. Inward-opening vent doors in the ram-air outlet header provide a means of pressure relief when the pressure differential in the air-conditioning bay reaches 0.4 psi. A water spray nozzle in the ram-air inlet is connected to a water drain in the condenser and to a second water drain downstream of the condenser. Air tapped from the air supply at the condenser drain and second drain sprays collected water into the heat exchanger inlet. Hot air from the engine bleed-air system flows to the primary heat exchanger. The airflow is partially cooled by heat transfer to ram air through the heat exchanger core and is then directed to the air cycle machine compressor inlet. From the compressor, the air flows through the secondary heat exchanger, where it is further cooled. The airflow is then directed, via the condenser, through the air cycle machine turbine.

Revision 2

21A-41

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONDITIONED AIR OUTLET TO CABIN COMPARTMENT

CONDENSED AIR OUTLET TO TURBINE INLET OF ACM

CONDITIONED AIR OUTLET TO CABIN COMPARTMENT

NOTE: NO. 1 SHOWN NO. 2 SIMILAR

MIXER

CONDENSER

WATER COLLECTOR RECIRCULATED AIR INLET WATER DRAIN

COLD AIR INLET (FROM TURBINE OUTLET OF ACM) BLEED-AIR INLET FROM SECONDARY SECTION OF DUAL HEAT EXCHANGER

Figure 21A-23. Condenser/Mixer

21A-42

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Condenser/Mixer

NOTES

The two condenser/mixers (Figure 21A-23) are all welded, aluminum alloy units consisting of a finned core and inlet and outlet headers. The units are bolted to structure in the rear fuselage, aft of the rear pressure dome. One end of the unit is the “mixer,’’ where cool air from its ACM, previously injected hot bypass bleed air, and recirculated cabin air mix together to achieve the temperature required in the cabin (No. 1 unit) or flight compartment (No. 2 unit). The other end of a unit is the “condenser,’’ where moisture is extracted from the turbine inlet air. Two drains, one fitted to the condenser outlet header and one downstream of the header, collect moisture extracted and route it back to its heat exchanger inlet, via a filter. Relatively cool conditioned air destined for the cabin or flight compartment passes through the finned core of either condenser. Warm air from the secondary heat exchanger passes over the finned core, where moisture in the air collects on the cool core. When water droplets form, they collect in water drains and are routed back to the heat exchanger ram-air inlet overboard.

Revision 2

21A-43

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TURBINE OUTLET (TO CONDENSER/ MIXER TURBINE HOUSING FAN INLET SCREEN

COMPRESSOR INLET (FROM PRIMARY HEAT EXCHANGER) COMPRESSOR HOUSING

HEAT EXCHANGER CONNECTING FLANGE FAN OUTLET DUCT

OIL FILL PLUG

OIL LEVEL SIGHT GAGE

FAN HOUSING

COMPRESSOR OUTLET (TO SECONDARY HEAT EXCHANGER)

R

Figure 21A-24. Air Cycle Machine

21A-44

Revision 2

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Air Cycle Machine The two air cycle machines (ACMs) each consist of a three-wheel assembly supported in a bearing cartridge and separate housings for compressor, turbine, and fan rotors (Figure 21A-24). The three housings are secured together by circular bolting flanges.

it is compressed and then delivered at a higher pressure and temperature to the secondary heat exchanger. The air is cooled in the secondary heat exchanger and is directed, via the c o n d e n s e r, t o t h e AC M t u r b i n e i n l e t . Expansion of the air across the turbine rotor reduces the pressure, with a corresponding drop in temperature. Air leaving the turbine outlet is routed to the condenser/mixer.

Compressor Housing The compressor housing is dome-shaped and flange-mounted on the turbine housing. The compressor inlet is located at the center of the domed housing. A diffuser assembly deflects the airflow leaving the centrifugal compressor toward an integral outlet duct in the housing.

Turbine Housing

The turbine extracts energy from the airflow as it reduces the pressure to just above cabin pressure. The major part of this energy is fed back to drive the compressor. With the axial flow fan rotor mounted on the same shaft, the remainder of the turbine energy drives the fan to ensure airflow through the heat exchangers in the absence of ram air.

NOTES

The turbine housing is the main structural member of the ACM supporting the compressor and fan housings on the end flanges and the bearing cartridge and three-wheel assembly in the bore. Part of the housing is machined to provide a sump that contains the oil required for lubrication of the ball bearing races and rotor shaft. The sump incorporates drain and filler plugs and an oil level sight gage.

Fan Housing The fan housing consists of two sections secured together at bolted flanges, with one section serving as the air inlet and the other providing the air outlet. The complete housing is flange-mounted to the turbine housing. The inlet section is machined to provide a circular, reverse-flow duct which directs ram air through the axial flow fan rotor. A wraparound wire mesh screen is riveted to the duct air intake to prevent the ingress of foreign objects and to protect the fan rotor. The fan rotates inside a steel shroud that is an integral part of the inlet section.

Operation Cooled air from the primary heat exchanger enters the compressor inlet of the ACM where

Revision 2

21A-45

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR CYCLE MACHINE

Figure 21A-25. Compressor Discharge Overtemperature Switch

21A-46

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Compressor Discharge Overtemperature Switch

NOTES

Each compressor discharge overtemperature switch (Figure 21A-25) is installed in a boss in the delivery duct from the ACM compressor. The switches consist of a normally open, single-pole thermal switch with a bimetallic element. Either switch closes when its respective compressor discharge temperature exceeds 207 ±5° C (405 ±10° F), with the following results: • The affected pack’s AIR COND PACK HOT caution light illuminates. • The affected pack’s pressure regulator and shutoff valve is energized closed, shutting off bleed air to the pack. • The recirculation air diverter valve (pre 1990 interior only) is positioned to shut off recirculation air to the affected pack and diverts all recirculation air to the operating pack. When the overheat condition clears, the overtemperature switch opens to restore the affected system’s operation, and the diverter valve is positioned to the center position to allow recirculation air to both packs.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

OPERATION (AIR-CONDITIONING PACKAGE) General Air to be cooled is reduced in temperature through either primary heat exchanger and is then compressed through its respective ACM to boost the temperature and pressure. The compressed air then passes through a secondary heat exchanger to further reduce the temperature, passes through a condenser where moisture is extracted, and enters the ACM turbine. The energy extracted by the turbine drops the air temperature to well below ambient and reduces the pressure to just above cabin ambient. At this point, bypassed bleed air is injected to achieve the desired temperature. The air then enters a mixing box, where it blends with recirculated cabin air.

Ram-Air Supply The ram-air supply consists of a baffle assembly and flexible duct, which connects the ram-air inlets on the dorsal fin to the conditioned-air supply duct. The baffle assembly allows the ram air either to enter the conditioned-air supply duct, if ram-air ventilation is required, or to spill over into the air-conditioning bay, where it is picked up by the ACM fans to cool the heat exchangers.

NOTES

Temperature of the air supply is regulated by opening and closing the pack temperature control valves to allow a proportion of bleed air to bypass the ACM and mix with the cool air discharged from the ACM. The valves are operated through mechanical linkage from a single actuator, which positions the valves in response to signals from an automatic temperature controller or from direct electrical inputs from a manually operated switch.

Pack Failure If a single air-conditioning pack fails, its pressure regulator and shutoff valve is energized closed, isolating the pack; the remaining pack then supplies conditioned air. If both air-conditioning packs fail, ram air will slightly pressurize the air-conditioning bay since neither ACM fan is operating. Ram air ceases to spill into the bay and enters the conditioned-air supply duct to ventilate the cabin and flight compartment.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TEMPERATURE CONTROL SYSTEM GENERAL The air-conditioning system temperature is controlled by two electrical subsystems, one for the cabin and one for the flight compartment. Each subsystem is controlled by independent sets of switches on the AIR CONDITIONING panel which can be positioned for manual or automatic operation. In the event of failure of one air-conditioning pack, air is supplied to the cabin and flight compartment by the remaining pack. A temperature gage on the AIR CONDITIONING panel monitors the temperature of the cabin, as well as the temperature in the cabin or flight compartment air ducts. The gage is controlled by a rotary threeposition switch (CAB DUCT—CABIN—FC DUCT).

duct temperature sensor, a duct temperaturesensing bulb, a duct overtemperature switch, a thermostatic switch, a zone temperature sensor, and a pack temperature control valve assembly (No. 2 pack), together with controls on the AIR CONDITIONING panel and associated relays. The system operates the same as the cabin system except that the signal from the temperature controller is sent to the No. 2 pack temperature control valve actuator to modulate the valves to achieve the desired flight compartment temperature.

NOTES

CABIN TEMPERATURE CONTROL SYSTEM The cabin temperature control system consists of a temperature controller, a duct temperature sensor, a duct temperature-sensing bulb, a duct overtemperature switch, a cabin temperature sensor, and a pack temperature control valve assembly (No. 1 pack), together with controls on the AIR CONDITIONING panel and associated relays. The temperature controller compares input signals from the cabin temperature selection made on a temperature selector on the AIR CONDITIONING panel and sends a signal to the appropriate pack valve actuator to modulate its valves to achieve the desired cabin temperature.

FLIGHT COMPARTMENT TEMPERATURE CONTROL SYSTEM The flight compartment temperature control system consists of a temperature controller, a

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN DUCT TEMPERATURE SENSOR

CABIN DUCT TEMPERATURESENSING BULB

CABIN DUCT OVERTEMPERATURE SWITCH

Figure 21A-26. Cabin Temperature Switch Locations DUCT TEMPERATURE DUCT SENSOR OVERTEMPERATURE TEMPERATURESWITCH SENSING BULB

Figure 21A-27. Flight Compartment Temperature Switches and Sensors

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Duct Temperature Sensors The duct temperature sensors (Figure 21A26 and 27) are negative-coefficient-type thermistors in which the resistance varies inversely with the temperature of the air flowing through the air-conditioning delivery ducts. Each sensor provides one of the signals to the associated temperature control. The cabin duct temperature sensors are located in the cabin supply duct in the rear fuselage aft of the rear pressure dome. The flight compartment duct temperature sensors are located in the flight compartment supply duct underneath the cabin floor.

Duct Overtemperature Switches Each duct overtemperature switch consists of a normally open, single-pole thermal switch with a bimetallic element. When the duct temperature exceeds 88° C (190° F), contacts in the switch close, and the appropriate cabin or flight compartment DUCT HOT caution light illuminates. Closing of either switch also energizes a duct overheat control relay (K3 and K4) to immediately switch the input to the appropriate pack bypass valve actuator to the manual COOL command. Contacts in the switch open when the duct air temperature falls to 82° C (180° F) to resume air-conditioning operation.

to temperature changes, to the TEMP gage on the air-conditioning panel. The gage reading allows for more precise temperature control when operating in the manual mode.

Zone Temperature Sensors The zone temperature sensors are negative coefficient, glass-bead-type thermistors, in which the resistance varies inversely with the cabin or flight compartment air temperature. The sensors provide one of the signals for automatic temperature control to the associated temperature controller. The cabin zone temperature sensor is located above the passenger service units in the fuselage, with a fan to blow air across it. The flight compartment zone temperature sensor is located behind the pilot in an exhaust air duct in the avionics bay.

NOTES

The duct overtemperature switch for each system is located adjacent to the duct temperature sensor in the cabin and flight compartment supply ducts.

Temperature-Sensing Bulb A temperature-sensing bulb is located in the cabin supply duct in the air-conditioning bay aft of the rear pressure bulkhead, and another is located in the flight compartment supply duct under the baggage compartment floor. They are electrical-resistant-type units which transmit variations in resistance, proportional

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 21A-28. Temperature Controllers

Figure 21A-29. Flight Attendant Temperature Control

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Temperature Controller

NOTES

The temperature controller for each system is located in the electrical equipment bay. Each controller consists of an aluminum alloy box which houses an electronic analog device using two-sided, printed circuit boards. A circuit in each controller is connected to its associated auto matic temperature control selector on the air-conditioning panel and to its associated duct and zone temperature sensors. A second circuit in each controller is connected to the applicable pack temperature control valve actuator and manual temperature control switch through auto/manual overheat relays. By comparing the input signals from the duct and zone sensors with the temperature selection made, the controller completes a circuit to operate the applicable pack temperature control valve actuator (No. 1 for cabin, No. 2 for flight compartment). The actuator, mechanically linked to valves, modulates the valves in response to the controller signals to supply conditioned air at the desired temperature. Power to operate the cabin temperature controller is supplied from the 28-VDC left secondary bus and for the flight compartment controller from the 28-VDC right secondary bus.

Flight Attendant Temperature Control There is a temperature control rotary switch on the forward flight attendant’s panel (Figure 21A-29). This switch gives the flight attendant total temperature control of the cabin conditioned air when the flight compartment airconditioning panel CABIN TEMP CONTROL rotary switch is selected to the full counterclockwise F/A position.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CAUTION LIGHT (AMBER) ILLUMINATED FLIGHT COMPARTMENT SUPPLY DUCT TEMPERATURE EXCEEDS 88° C.

CAUTION LIGHT (AMBER) ILLUMINATED FLIGHT COMPARTMENT PACK COMPRESSOR DISCHARGE TEMPERATURE EXCEEDS 207° C (405 °F).

PACK TEMPERATURE CONTROL VALVES ARE DRIVEN TO MAXIIMUM COOLING POSITION WHILE OVERTEMPERATURE PERSISTS.

5

6

7

A

FLT COMPT PACK HOT

B

#1 BLEED HOT

C

FLT COMPT DUCT HOT

D

CABIN DUCT HOT

E

CABIN PACK HOT

PRESSURE REGULATOR AND SHUTOFF VALVE IS ENERGIZED CLOSED TO CUT OFF BLEED-AIR SUPPLY.

#2 BLEED HOT

CAUTION LIGHT (AMBER) ILLUMINATED ASSOCIATED NACELLE SHUTOFF VALVE SUPPLY DUCT TEMPERATURE EXCEEDS 290° C (550° F) OR DUCT OVER PRESSURE OF 75 PSI.

F

CABIN PRESS

W

4

5

6

7

CAUTION LIGHT (AMBER) ILLUMINATED CABIN PACK COMPRESSOR DISCHARGE TEMPERATURE EXCEEDS 207° C.

WARNING LIGHT (RED) ILLUMINATED CABIN ALTITUDE EXCEEDS 10,000 FEET. THE LIGHT GOES OUT WHEN CABIN ALTITUDE DESCENDS BELOW 10,000 FEET.

PRESSURE REGULATOR AND SHUTOFF VALVE IS ENERGIZED CLOSED TO CUT OFF BLEED-AIR SUPPLY. CAUTION LIGHT (AMBER) ILLUMINATED CABIN SUPPLY DUCT TEMPERATURE EXCEEDS 88° C. PACK TEMPERATURE CONTROL VALVES ARE DRIVEN TO MAXIMUM COOLING POSITION WHILE OVERTEMPERATURE PERSISTS.

Figure 21A-30. Caution and Warning Lights

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DUCT OVERHEAT SYSTEM

NOTES

If the cabin or flight compartment supply duct air temperature exceeds 88° C (190° F), a duct overtemperature switch closes, with the following results: • An electrical circuit is completed through the switch to the master caution control box, and a DUCT HOT caution light (Figure 21A-30) illuminates. • The duct overheat relay K3 or K4 is energized, initiating the following functions: • The auto/manual relay K1 or K2 is energized to isolate the automatic hot and cold commands from the cabin or flight compartment temperature controller to the associated pack temperature control valve actuator. • An electrical circuit is completed to provide a direct “cool’’ signal to the associated pack valve actuator. This signal causes the actuator to reposition its valves to operate the associated pack in the full cold mode. When the duct temperature falls to 82° C (180° F), the affected overtemperature switch opens to resume air-conditioning operation.

NOTE When the APU (if installed) is used as the source of bleed air, a duct overheat condition will result in the APU bleed valve closing and cutting off the bleed air.

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21A-55

21A-56 CABIN TEMPERATURE CONTROLLER

28V DC R108

FLIGHT COMPARTMENT TEMPERATURE CONTROLLER

R101

P4

I

J G H A B C D E

F

J2 C D E F

D

C

A

B

E D3

B3 B1 A3 K1 A1 AUTO MANUEL RELAY

3

1

10K HOT

CABIN ZONE TEMP SENSOR

FLIGHT ATTENDANT CABIN TEMP SELECTOR

D3 A B C D E F

F D2

B

C2

A

B2

(CCW) 3

COLD

2

2

R4

D1 C3 C1

R3

1

C

A2

D

X1 X2

NO.1 PACK VALVE ACTUATOR MAKES CONTINUITY BETWEEN PINS H & E WHEN PACK HOT BYPASS VALVE CLOSED

GND CASE GND CCW HI COLD CCW HI HOT CCW LOW

F/COMP TEMPERATURE SELECTOR

EMI FILTER

S1

HOT

S2

COLD

1.5A START UP .5A STEADY STATE S1 SWITCHES TO OPPOSITE CONTACT AT CCW LIMIT S2 SWITCHES TO OPPOSITE CONTACT AT CW LIMIT

F/COMP ZONE TEMP SENSOR

CW LOW

C2

RE

1 R1

10K

M

3

2 3

P7 A B C D

'Q'

D2

'R'

D1 C3 C1

}

B1 A3 A1 X1

P19 A

1 2 3

4 5 6

B C D CABIN TEMP AUTO

CABIN TEMP MAIN 5

28V DC L SEC

COOL 3

28V DC R SEC

3

WARM OFF

CABIN TEMP MANUAL CONTROL SWITCH S1

C2 LOGIC CIRCUIT

F/COMP DUCT TEMP SENSOR

B2

B3 2

P16

34K @ 71° F

P1 H G

X2

(CW)

CABIN DUCT TEMP SENSOR

P5

X1

CABIN TEMPERATURE SELECTOR

34K @ 71° F

B

A2

A1

COLD

A

K13

A3

HOT

G H

DUCT OVERTEMP SWITCHES (CLOSE AT TEMP > 190° F (88° C) ) CABIN F/C

A2

X2

K3 DUCT OVERHEAT RELAY

P24

A

B

A

B

'S'

P9 A

CABIN DUCT HOT

FLT COMP HOT

34K AT 71° F

B C D 1 2 3

4

5 6

P12 F/COMP TEMP AUTO

CR2 CR3

COOL

F/COMP TEMP M/CONT 5

28V DC R SEC

5 28V DC L SEC

WARM

3 OFF FLT/COMPT TEMP MANUAL CONTROL SWITCH S2 'T' 'U' 'V' 'L' 'M' 'N'

Figure 21A-31. Temperature Control Electrical Schematic—Series 300 (Sheet 1 of 2)

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

I J

P3

Canada Ltd.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TEMPERATURE CONTROL OPERATION

NOTES

NOTE The following is a condensation of a detailed text in Chapter/Section 21-60-00 of the Series 300 Maintenance Manual and relates to MSM Chapter 21A.

Automatic Temperature Control Cabin In the automatic mode, the left PACKS switch on the AIR CONDITIONING panel is positioned to AUTO, and the CABIN selector is rotated toward COOL or WARM, as required (Figure 21A-31). The cabin temperature controller transmits electrical pulses to the No. 1 pack temperature control valve for regulation of airflow to the desired temperature.

Flight Compartment In the automatic mode, the right PACKS switch is positioned to AUTO, and the FLT COMP temperature control selector is rotated toward COOL or WARM, as desired. The flight compartment temperature controller transmits electrical pulses to the No. 2 pack temperature control valve for regulation of airflow to the flight compartment at the desired temperature.

Manual Temperature Control Cabin In the manual mode, the left PACKS switch is set to MAN, and the desired temperature is selected with the CABIN manual temperature control selector switch positioned toward COOL or WARM, as desired. Hot or cold command signals from the cabin temperature controller are isolated from the No. 1 pack temperature control valve actuator. Control signals are transmitted directly to the control valve actuator for positioning in accordance with the manually selected temperature.

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21A-57

21A-58 MAN

H G

A

B2 B3 B1

AUTO/ MAN RELAY K2

C

A2

S1

EMI FILTER

3

X1 K14 X2

CR3

X1 CR4

X2

D3

A3

C2

C3 C1

X1 X2

A

C

B1 A3 A1

A2

X2

K4 DUCT OVERHEAT RELAY

P17

DIVERTER VALVE ROTARY ACTUATOR

S7 C

A X1

D1

X2

D3

A2

A3 A/C A3 A/C OVERTEMP OVERTEMP A1 A1 SWITCH #1 SWITCH #2

X1

1 HI

CABIN DUCT

D3

K9 X2

3 LO

K10

A B CR9 P23

P22 A B

2 BOTH

3 CABIN AIR SUPPLY DIVERTER VALVE

FLT COMPT. PACK HOT

CABIN PACK HOT

C2

A1

B C

K12 PRESSURE K11 T/D ON OPERATE REGULATOR AND (20 SECS) SHUT-OFF VALVE (PACK)

B

B3

D1

C3

P20

A2 A1

D2

X1

'V' 'L' 'M' 'N'

MAN

A1

M

S2

B2

'T' 'U'

1 OFF

4

A2

A3

'S'

S8

A3

X2

P21 'R'

MAN

X1

K15

A2 'Q'

2

S9 5

1.5A START UP SA STEADY STATE

CW LOW

D

AUTO

1 3

CCW HI COLD CW HI HOT CCW LOW

B

C2

3

A1

CASE GND

F D2

D1 C3 C1

2

GND

E D3

A3

A2

AUTO

P2 P6

6

CR5

B

5

S1 SWITCHES TO OPPOSITE CONTACT AT CCW LIMIT S2 SWITCHES TO OPPOSITE CONTACT AT CW LIMIT

B3 B2 A2

A3

B1 A2

A1 PRESSURE REGULATOR AND SHUT-OFF VALVE (PACK 2)

X1 X2

C1 B3

B2 B1

A3 A1

X1

P18

X2

K7

K16

D B

DIVERTER VALVE ROTARY ACTUATOR (CCW)

A2 (CW)

CLOSES 405° ± 10F

M

A

X1 X2

C E NOTES: 1

INDENT CODE IS 2161-UNLESS OTHERWISE SPECIFIED.

2

FOR CONTINUATION SEE CABIN DUCT TEMPERATURE INDICATOR 21-62-11

3

PARTS LOCATED ON AIRCONDITIONING PANEL IN COCKPIT OVERHEAD CONSOLE

LOGIC CIRCUIT

Canada Ltd.

Revision 2

Figure 21A-31. Temperature Control Electrical Schematic—Series 300 (Sheet 2 of 2)

RECIRCULATION DIVERTER VALVE

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

D C A

4 OFF

NO.2 PACK VALVE ACTUATOR

FLIGHT COMPARTMENT TEMPERATURE CONTROLLER

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Flight Compartment

NOTES

In manual mode, the right PACKS switch is set to MAN, and the desired temperature is selected by adjusting the FLT COMP manual temperature control selector switch toward COOL or WARM. The hot and cold command signals are disconnected from the flight compartment temperature controller. Circuits are activated between the FLT COMP temper ature control selector and the No. 2 pack temperature control valve actuator. Positioning the selector toward COOL or WARM directly positions the temperature control valve for the desired temperature.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN GRILLE GALLERY

DADO TO LAVATORY PANEL AND FLIGHT ATTENDANT'S GASPER

GASPER GALLERY

DUCT TEMPERATURE SENSOR

RISER

DUCT TEMPERATURE SENSING BULB

DUCT OVERTEMPERATURE SWITCH

CABIN AIR

Figure 21A-32. Cabin Air Distribution Schematic

21A-60

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONDITIONED AIR DISTRIBUTION

NOTES

CABIN AIR DISTRIBUTION Conditioned air for the cabin is supplied by the No. 1 (rear) air-conditioning pack (Figure 21A-33). The air enters the cabin via the center of the rear pressure dome and is routed below the baggage compartment and cabin to outlets in the dado panel above the floor. Air is also ducted from the under floor cabin duct to upper outlets in the passenger service units (PSU) in the cabin via six left and five right sidewall risers. In addition to normal cabin air conditioning, manually controllable air outlets (gaspers) are provided at each seat location, in the center ceiling above the forward flight attendant’s station, the lavatory, and the flight compartment. Except for those in the flight compartment, all gaspers receive conditioned air from the cabin air supply duct via takeoff ducts from the same sidewall risers as the upper cabin (PSU) outlets. Approximately 60% of cabin air is exhausted through slots, located between the side and dado panels of the mid and aft cabin, passing under the floor to be exhausted by the normal outflow valves. The other 40% of cabin air is exhausted through the air recirculation system. The cabin air delivery duct from the No. 1 pack incorporates a duct temperature sensor and a duct overtemperature switch, which are part of the temperature control system. A cabin duct temperature-sensing bulb in the cabin supply duct downstream of the duct overtemperature switch provides signals to a cabin duct temperature indicator on the AIR CONDITIONING panel in the flight compartment.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FORWARD PRESSURE BULKHEAD

REAR PRESSURE BULKHEAD

SIDE WINDOW DEMIST OUTLETS

FLOW ADJUST LEVERS

CABIN DUCT OVERTEMPERATURE SWITCH TO LARGE FORWARD SIDE CONSOLE GASPERS

FIXED OUTLETS IN FOOTWELL

SIDE CONSOLE ADJUSTABLE OUTLETS

TO GASPERS BELOW WINDSHIELD CORNER PILLARS

DUCT TEMPERATURE SENSOR

CABIN SUPPLY

FLOW ADJUST LEVERS

ZONE TEMPERATURE SENSOR

SIDE WINDOW DEMIST OUTLETS

T RR

Figure 21A-33. Flight Compartment Air Distribution

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT AIR DISTRIBUTION (1990 INTERIOR)

NOTES

Description The supply of conditioned air for the flight compartment, consisting of approximately 30% of the air discharged from the air-conditioning pack, enters the cabin at the center of the rear pressure dome and is routed to the flight compartment via a duct under the cabin floor (Figure 21A-33). A wye duct splits the airflow into two individual but identical air distribution subsystems, one for the pilot and one for the copilot. Each of the two distribution subsystems consists of a side window demist twin outlet, a fixed outlet in the footwell, and an adjustable large outlet in the side console. Flow adjust levers on the sill below the side window are connected by cables which operate movable vanes in the supply ducts to permit adjustment of airflow. The pilot and copilot are each provided with one large and one small manually controllable (gasper) air outlet. The large gaspers are located at the forward end of each side console; the small gaspers are located below the windshield pillars on both sides of the flight compartment. Air is exhausted from the flight compartment under the floor to the recirculation duct. From there the air is drawn back to the pack condenser/mixer by the recirculation fan. The delivery duct from the pack incorporates a duct temperature sensor and an overtemperature switch, which are part of the temperature control system.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

LAVATORY GASPER

FLIGHT ATTENDANT'S GASPER

GASPER GALLERY

REAR PRESSURE BULKHEAD

FLIGHT COMPARTMENT CABIN RESTRICTOR RISERS

LOWER CABIN AIR FLIGHT COMPARTMENT AIR

LAVATORY TEC

Figure 21A-34. Gasper Systems Schematic

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

GASPER AIR SYSTEM (1990 INTERIOR)

NOTES

The cabin gasper air system comprises two separate and independent systems, one each for the cabin and flight compartment (Figure 21A-34). The system is supplied with conditioned air from the lower cabin conditioned air supply duct, via takeoffs from six sidewall risers (three on each side of the fuselage). The risers route air to the gasper galleries that extend above the passenger service unit (PSU) panels on each side of the cabin. From there the air is supplied to individual controllable outlets at each passenger seat. The risers also supply conditioned air to the cabin upper air outlet grills. The upper cabin grills require less airflow than the gaspers; therefore, a restrictor is installed in each riser between the gasper takeoff and upper grille gallery. An additional single riser supplies cabin duct air to a single lavatory gasper and a single flight attendant’s station ceiling gasper. The flight compartment gasper system is supplied with conditioned air from the flight compartment conditioned-air supply duct. Two individual controllable outlets are supplied for the pilot and copilot. A large outlet is located at the forward end of the pilot’s and copilot’s side consoles, while a smaller outlet is located below the pilot’s and copilot’s windshield corner pillars.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT

AVIONICS COMPARTMENT

OUTFLOW VALVES

CABIN RECIRC AIR EXTRACTION DUCTS

BAGGAGE COMPARTMENT

AIR EXHAUST SLOT BETWEEN SIDE AND DADO PANELS

NO. 2 RECIRC FAN (FLT COMPT)

RECIRCULATION DUCT

RECIRCULATION FAN NO. 2 B A

P2 C D

2123-K2 FAN CONTROL RELAY

CHECK VALVES

A B

NO. 1 RECIRC FAN (FLT COMPT)

RECIRCULATION FAN NO. 1 C D P1

2123-K1 FAN CONTROL CR1 RELAY

CR2

2123-S2 OFF

2123-S1 OFF RECIRC FLT COMP

RECIRC FAN 5 CONT 28 VDC RIGHT SEC BUS

50 RECIRC FAN PWR 28 VDC RIGHT SEC BUS

RECIRC CABIN

RECIRC 50 FAN PWR 28 VDC LEFT SEC BUS

RECIRC FAN 5 CONT 28 VDC LEFT SEC BUS

Figure 21A-35. Air Recirculation System and Electrical Schematic

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR RECIRCULATION SYSTEM (1990 INTERIOR) Description The recirculation system routes exhaust air from under the flight compartment floor and the cabin, via air extraction ducts located behind the top of the dado panels adjacent to the forward passenger seat rows (Figure 21A36). From there, the air is drawn back through the recirculation air duct to the two pack condenser/mixers by their respective recirculation fans. A check valve is installed between each recirculation fan and its respective pack.

20 volts. Occasionally, an overcurrent occurs requiring that the fan be switched off for 20 seconds prior to restart. Overheat also causes the fan to shut down, but it automatically restarts when the overheat condition clears.

NOTES

The fans are controlled, through two fan control relays, by two switches on the air-conditioning panel. The left switch is marked OFF—RECIRC CABIN and controls the rear No. 1 pack fan. The right switch is marked OFF-RECIRC FLT COMP and controls the No. 2 pack fan.

Operation Selecting the left switch to RECIRC CABIN completes a circuit to energize fan control relay 2123-K1, completing a circuit from the left 28-VDC secondary bus to the No. 1 recirculation fan. Selecting the right switch to RECIRC FLT COMP completes a circuit to energize fan control relay 2123-K2. The relay then completes a circuit from the right 28-VDC secondary bus to the No. 2 recirculation fan.

NOTE The recirculation fans incorporate protective features for undervoltage, overcurrent and overheat. The symptoms are those of a fan failure. Recir culation fans automatically shut down when exposed to transient voltages of less than 22 volts for periods in excess of 0.10 second. They automatically restart between 18 and

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EQUIPMENT COOLING FAN

LOW SPEED WARNING DEVICE

THERMOSTATIC SWITCH

AIR EXTRACTION DUCT

FLT COMP

AVIONICS COMPT

CABIN

FLT COMPT SENSOR

AVIONICS FAN

LOW SPEED WARNING DEVICE P1

DIODE WARNING LIGHT DS1

F

E

A

B C

J2

(20 SEC DELAY ON OUTPUT PIN F)

B2 A2

P2 C B A

FAN

B3 B1 A3 A1

A3

X1

A1

X2

A2

K2

X1 X2

K1

S1 CLOSES AT > 35°C

COOLING FAN

10 28V DC L MAIN BUS P/O LEFT AVIONICS CIRCUIT BREAKER PANEL

NOTE: IDENT CODE IS 2126 UNLESS OTHERWISE INDICATED

Figure 21A-36. Avionics Compartment Cooling Fan

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AVIONICS COMPARTMENT COOLING (1990 INTERIOR) Avionics compartment cooling is accomplished by an avionics fan located under the cabin floor (Figure 21A-36). The fan draws air off each of the equipment shelves, via air extraction ducts, and discharges the air under the cabin floor. The brushless, thermostatically controlled fan is operational whenever DC power is on. In the event of fan failure, cooling by natural convection is adequate for interim operation. Fan failure is indicated by a light on the flight attendant’s panel.

NOTE A 20-second delay is incorporated in the low speed warning device on output pin F. This prevents a false fan failure indication while the fan is spooling up to operational speed after initial start.

NOTES

OPERATION Normal The avionics cooling fan operates when the ambient temperature at the avionics rack exceeds 35° C. (As detected by the thermostatic switch mounted on the upper avionics shelf). In this condition, switch S1 closes and relay K1 is energized. 28 VDC is applied through the fan circuit breaker and contacts of relay K1 and K2 to energize the fan and provide power to the low speed warning device. Normal operating speed for the fan is 9000 rpm.

Fan Failure Speed of the fan is monitored at pin C of the low speed warning device by an output from the fan. If the speed of the fan drops to 6,000 rpm or less, and remains at the reduced speed for more than 20 seconds, the low speed warning device provides an output which energizes relay K2. Contacts of K2 disconnect power from the fan; K2 also maintains power to the low speed warning device independent of relay K1. The low speed warning device also turns on the diode warning light DS1 to indicate fan failure.

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CHAPTER 21B (SERIES 200) AIR CONDITIONING CONTENTS Page INTRODUCTION .............................................................................................................. 21B-1 BLEED-AIR SYSTEM—Series 200.................................................................................. 21B-3 General........................................................................................................................ 21B-3 Components Description and Operation..................................................................... 21B-5 System Operation...................................................................................................... 21B-21 AIR-CONDITIONING SYSTEM—SERIES 200 ........................................................... 21B-25 General...................................................................................................................... 21B-25 Components Description and Operation................................................................... 21B-29 Operation ................................................................................................................. 21B-41 TEMPERATURE CONTROL SYSTEM......................................................................... 21B-43 General...................................................................................................................... 21B-43 Operation .................................................................................................................. 21B-45 CONDITIONED AIR DISTRIBUTION.......................................................................... 21B-49 General...................................................................................................................... 21B-49 Cabin Air Distribution .............................................................................................. 21B-51 Flight Compartment Air Distribution ....................................................................... 21B-53 Gasper Air System.................................................................................................... 21B-53 Air Recirculation System.......................................................................................... 21B-55 Avionics Compartment Cooling ............................................................................... 21B-57 Operation .................................................................................................................. 21B-57

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PRESSURIZATION SYSTEM—SERIES 200................................................................ 21B-59 General...................................................................................................................... 21B-59 Components Description and Operation................................................................... 21B-61 Operation .................................................................................................................. 21B-65

21B-ii

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ILLUSTRATIONS Figure

Title

Page

21B-1

Bleed-Air System Schematic............................................................................... 21B-2

21B-2

High-Pressure Switches ....................................................................................... 21B-4

21B-3

High-Pressure Bleed Shutoff Valve ..................................................................... 21B-6

21B-4

Handling Bleed Valve .......................................................................................... 21B-8

21B-5

Precooler Installation......................................................................................... 21B-10

21B-6

Low-Pressure Check Valve................................................................................ 21B-12

21B-7

Nacelle Shutoff Valve ........................................................................................ 21B-14

21B-8

Bleed-Air Overtemperature Switch ................................................................... 21B-14

21B-9

Bleed-Air Overpressure Switch......................................................................... 21B-15

21B-10 Pressure Regulator Valve................................................................................... 21B-16 21B-11 Pressure Regulator and Shutoff Valve Schematic (Normal Regulated Flow) ... 21B-18 21B-12 Bleed-Air System Electrical Schematic—Series 200........................................ 21B-20 21B-13 Air-Conditioning System Controls.................................................................... 21B-24 21B-14 Air-Conditioning Pack Schematic ..................................................................... 21B-26 21B-15 Temperature Trim Valves................................................................................... 21B-28 21B-16 Pack Temperature Control Valves ..................................................................... 21B-30 21B-17 Heat Exchanger.................................................................................................. 21B-32 21B-18 Condenser/Mixer ............................................................................................... 21B-34 21B-19 Air Cycle Machine ............................................................................................ 21B-36 21B-20 Compressor Discharge Overtemperature Switch............................................... 21B-38 21B-21 Temperature Controllers .................................................................................... 21B-38 21B-22 Temperature Sensors and Switches ................................................................... 21B-40 21B-23 Temperature Control System ............................................................................. 21B-42

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21B-24 Cabin Temperature Bias Control ....................................................................... 21B-44 21B-25 Conditioned Air Distribution............................................................................. 21B-48 21B-26 Cabin Air Distribution Schematic ..................................................................... 21B-50 21B-27 Flight Compartment Air Distribution ................................................................ 21B-52 21B-28 Gasper Systems Schematic................................................................................ 21B-52 21B-29 Air Recirculation System Schematic................................................................. 21B-54 21B-30 Avionics Compartment Cooling Fan ................................................................. 21B-56 21B-31 Pressurized Areas .............................................................................................. 21B-58 21B-32 Pressurization Controls and Indicators .............................................................. 21B-58 21B-33 Normal Outflow Valve—Series 200.................................................................. 21B-60 21B-34 Selector Panel .................................................................................................... 21B-62 21B-35 Indicator Panel................................................................................................... 21B-64 21B-36 Pressurization System Schematic ...................................................................... 21B-66 21B-37 Pressurization Envelope..................................................................................... 21B-68 21B-38 Forward Dump Manual Selector ....................................................................... 21B-68 21B-39 Pressurization Control—Electrical Schematic .................................................. 21B-70

21B-iv

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TABLE Table 21B-1

Revision 3

Title

Page

Pressurization Control Settings ......................................................................... 21B-64

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CHAPTER 21B AIR CONDITIONING (SERIES 200)

INTRODUCTION This chapter, though titled “Air Conditioning,’’ deals with the environmental systems of the Dash 8, including bleed air, air conditioning, and pressurization. Information is included from Chapter 21, “Air Conditioning’’ and Chapter 36, “Pneumatics,’’ of the 200 Series Maintenance Manual. The material in this chapter is oriented toward the line mechanic. All values expressed throughout this chapter, such as for pressure, temperature, flow rates, and time, are used only for their illustrative meanings. Actual values may differ and must be obtained from the pertinent sections of the Maintenance Manual.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

1A

1B

3

6

2 HP 2 NO. 2 ENGINE

5

LP

TO DEICING SYSTEM

4 7 11

AIR CONDITIONING OFF

1

RECIRC

40 20 °C 0 DUCT TEMP

12 BLEED

2

8 9

60

#2 BLEED HOT

80 100

MIN

MAX

10

BLEED

13

MAN AUTO COOL

WARM

PACKS

COOL

WARM

10 F/A CABIN

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

TEMP CONTROL

TO AIR-CONDITIONING PACK

FLT COMP

HIGH-PRESSURE SWITCHES CHOKING VENTURI HIGH-PRESSURE SHUTOFF VALVE LOW-PRESSURE CHECK VALVE HANDLING BLEED VALVE (HBOV) PRECOOLER NACELLE SHUTOFF VALVE BLEED OVERPRESSURE SWITCH OVERTEMPERATURE SWITCH WING ISOLATION CHECK VALVE SERVO AIR FILTER FLOW CONTROL SERVO PRESSURE REGULATOR VALVE

FROM NO. 1 ENGINE BLEED AIR SYSTEM (SIMILAR TO NO. 2 SYSTEM)

Figure 21B-1. Bleed-Air System Schematic

21B-2

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BLEED-AIR SYSTEM— SERIES 200

NOTES

GENERAL The bleed-air system in each nacelle consists of an optional precooler with a removable bypass section, two high-pressure switches, a high-pressure shutoff valve, high-pressure choking venturi, low-pressure check valve, nacelle shutoff valve, and duct overpressure and overtemperature switches (Figure 21B-1). The high- and low-pressure engine bleeds join downstream. Wing isolation check valves route bleed air to an air-conditioning pack in the aft fuselage. The pressure regulator valve is controlled by a single flow control servo unit. Each nacelle bleed-air system is controlled by a two-position bleed switch on the AIR CONDITIONING panel (Figure 21A-1). The BLEED flow control knob adjusts the pressure regulator, controlling the flow rate. The engines have a high flow and pressure ratio. To achieve the higher first-stage pressure at rated power settings, it is necessary to spill excess low-pressure air with the engine at idle. This is achieved with a bleed-off valve (HBOV) at the low-pressure bleed port. The valve, installed by the engine manufacturer, is actuated by a torque motor controlled by the engine electronic control (EEC). It is open during ground idle and ramped closed as engine rpm increases. The valve also opens at flight idle near sea level, but not at operating altitude (25,000 feet) except during engine restart.

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NOTE: SWITCHES IN BOTH LOCATIONS ARE IDENTICAL EXCEPT FOR PRESSURE SETTINGS.

Figure 21B-2. High-Pressure Switches

21B-4

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COMPONENTS DESCRIPTION AND OPERATION

NOTES

High-Pressure Switches Two high-pressure switches (Figure 21B-2), located on the wing front spar outboard of the nacelle, are connected to a tapping in the outlet duct from the engine high-pressure bleed port. Each switch is a normally closed pneumatic pressure-sensing switch containing diaphragm-operated electrical contacts which are connected into the circuit to the high-pressure shutoff valve when the BLEED switch is selected to BLEED. The No. 1 switch is set at 65 psig; the No. 2 switch is set at 55 psig.

Operation In ground taxi mode with weight on wheels (WOW), the No. 1 switch (65 psig) is activated, allowing the airplane bleed air to remain on high pressure through taxiing. A sensed pressure of 65 psig increasing moves the diaphragm to open a set of contacts in the switch; the contacts close when the sensed pressure drops to 65 psig. This switch is also activated when the deice AIRFRAME AUTO switch is selected to SLOW or FAST. In the flight mode (no WOW signal), the only difference from the ground mode is that the No. 2 (55 psig) switch is activated instead of the No. 1 switch. However, when the deice AIRFRAME AUTO selector is positioned to SLOW or FAST, bleed air will be controlled by the No. 1 switch.

High-Pressure Bleed-Air Venturi A choking venturi (Figure 21B-1) installed in the HP bleed port restricts bleed-air flow to a maximum of 10%. This restriction prevents engine bleed from reaching damaging proportions, such as both air-conditioning packs operating from a single engine or in the event of a duct rupture.

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SOLENOID ACTUATOR HOUSING

ELECTRICAL CONNECTOR ACTUATOR COVER

ACCESS TO FILTER

FLOW CONTROL VALVE

SOLENOID SHOWN ENERGIZED FROM PRESSURE SWITCH

A B C

SOLENOID SHOWN DEENERGIZED FROM PRESSURE SWITCH

A B C

CONNECTOR

CONNECTOR VALVE OPEN

FILTER

VALVE OPEN

FILTER DIRECTION OF FLOW

DIRECTION OF FLOW

TWO-INCH DISC

Figure 21B-3. High-Pressure Bleed Shutoff Valve

21B-6

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High-Pressure Bleed Shutoff Valve

NOTES

The high-pressure (HP) bleed shutoff valve (Figure 21B-3) is line-mounted in the hot air duct from the high-pressure bleed port on the engine. It is a normally closed, solenoid-controlled pneumatically operated valve used to control the flow of HP bleed air from the engine. The valve consists of a solenoid and a pneumatic actuator which is mechanically linked to a butterfly valve. When the solenoid is energized, servo pressure from upstream is applied to the pneumatic actuator to open the butterfly. To ensure correct installation of the valve in the system, a flow direction arrow is provided on the valve body.

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21B-7

21B-8

DUCT TUBE

COVER

RESTRICTOR P2.4

P2.4

TO EEC

SERVO CONTROLLED P2.4 BLEED

PISTON SERVO VALVE

ELECTRICAL HARNESS CONNECTION

P2.5

OPEN POSITION

Figure 21B-4. Handling Bleed Valve

P2.5

CLOSED POSITION

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Handling Bleed Valve Control

NOTES

The handling bleed valves (HBV) (Figure 21B-4), located on each engine, are connected with and control the outflow of the associated engine P 2.5 bleed air overboard through a precooler in the bleed-air system The HBV is used to prevent compressor surge and stall and to reduce noise level during steady operation of the engine. It consists of a housing, a piston in a ported sleeve, a cover, a servo valve, and a manifold with a restrictor. The HBV piston is controlled by P 2.5 air on its underside and by P 2.4 air on the opposite side. P 2.4 air is ducted from a diffuser pipe through the restrictor and manifold, then into the servo valve. On a signal from the EEC, the flapper nozzle in the servo valve opens and vents P 2.4 air. P 2.4 air in the HBV undergoes a pressure drop, which allows P 2.5 air to seat the piston. The valve is then open, and P 2.5 air vented into an airframe tube is ducted to the airframe precooler (to reduce P3 bleed temperature) and is then vented overboard. On a signal from the EEC, the flapper nozzle in the servo valve closes; P 2.4 air stops venting, and its pressure stabilizes to seat the piston against P 2.5 air. The piston stops P 2.5 air from entering the valve and venting. The valves are controlled by the corresponding engine ECU through a potentiometer on the power lever layshaft.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FWD

PRECOOLER UNIVERSAL JOINT

HP SHUTOFF VALVE

TAP OFF TO HP SWITCHES

HP VENTURI

RR3

Figure 21B-5. Precooler Installation

21B-10

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Precooler

NOTES

The precooler (Figure 21B-5) is a steel heat exchanger mounted above the engine in the high-pressure (HP) bleed-air supply line. Its purpose is to cool the hot HP engine bleed air to an acceptable temperature for the aluminum heat exchangers of the two rear fuselage air-conditioning packs. The cooling effect is achieved by the handling bleed-off (HBOV) air flowing across the precooler and out the zone 2 exhaust louver on the top of the engine nacelle.

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HOUSING ASSEMBLY

SEAT

OPEN POSITION

PISTON

CLOSED POSITION

HANDLING BLEED VALVE

INTERCOMPRESSOR CASE (REF) P2.5

Figure 21B-6. Low-Pressure Check Valve

21B-12

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Low-Pressure Check Valve

NOTES

The low-pressure check valve (Figure 21B-6) is mounted in the low-pressure (LP) bleed port. Its prime function is to isolate LP bleed air from the bleed-air system when the HP bleed system is operating. The LP bleed port does not have a venturi like the HP system, but the check valve has a restrictor built in to perform the same function. The valve limits LP bleed to a maximum of 10% of bleed flow.

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Figure 21B-7. Nacelle Shutoff Valve

Figure 21B-8. Bleed-Air Overtemperature Switch

21B-14

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Nacelle Shutoff Valve

Bleed-Air Overpressure Switch

The nacelle shutoff valve is installed in the outlet duct of the nacelle bleed-air system (Figure 21B-7). The function of the valve is to shut off bleed air to the air-conditioning pack when they are not selected and when the engine high-pressure bleed port overpressure or overtemperature limits are exceeded.

The wing bleed overpressure switch is installed in the bleed-air delivery duct, downstream from the nacelle shutoff valve (Figure 21B-9). If bleed-air pressure exceeds 75 psig, the switch completes a circuit to close the nacelle shutoff valve, isolating the affected wing and illuminating the appropriate BLEED HOT caution light.

Overtemperature Switch

NOTES

The overtemperature switch (Figure 21B-8) is installed in the bleed-air delivery duct downstream of the nacelle shutoff valve. It consists of a normally open thermal switch with a bimetallic element and closes when air temperature in the duct exceeds 550 +10° F (287 +5° C) and illuminates the appropriate BLEED HOT caution light.

Figure 21B-9. Bleed-Air Overpressure Switch

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CONTROL COVER

ACCESS TO FILTER

SOLENOID (DISCONNECTED)

ACTUATOR AND TUBE CONTROL HOUSING

PRESSURE TEST PORT

ACTUATOR HOUSING

FLOW CONTROL VALVE

Figure 21B-10. Pressure Regulator Valve

21B-16

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Pressure Regulator Valve (Figure 21B-10)

NOTES

The pressure regulator is installed in the bleedair duct in the rear fuselage aft of the rear pressure bulkhead. The purpose of the valve is to regulate the flow of bleed air into the system. It consists of a pneumatic actuator mechanically connected to a butterfly valve, and a pressure-sensing control to maintain a schedule of bleed airflow.

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BLEED FLOW CONTROL ELECTROPNEUMATIC SERVO BLEED MAX MIN

TORQUE MOTOR FILTER

28 VDC L&R SEC MOD 8/1820

18 PSI FROM DEICE SYSTEM

PRESSURE SENSING CONTROL

SERVOACTUATOR

STABILIZER UNIT

ANEROID RESTRICTOR

AMB SUMMING BAR

PNEUMATIC ACTUATOR

SENSING ACTUATOR TEST PORT CLOSE FILTER

LEGEND

AMB

BLEED AIR REGULATED AIR REGULATED AIR (18 PSI) AMBIENT BUTTERFLY VALVE (OPEN)

Figure 21B-11. Pressure Regulator Shutoff Valve Schematic (Normal Regulated Flow)

21B-18

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Pneumatic Actuator Operation Pressure for valve operation is from upstream of the valve and is routed to both sides of the pneumatic actuator piston (Figure 21B-11). The close side is spring-assisted; the open side has a larger area to overcome the combined force of the spring and air pressure. The amount of pressure to the open side is controlled by the pressure-sensing control, which consists of four pneumatic inputs to a summing bar that controls pressure venting to establish open side pressure. These inputs are: • An actuator sensing downstream pressure • A stabilizer unit controlled by a restricted line • An aneroid that senses pressure altitude pressure changes

the sensing actuator returning to normal and less air being vented. The stabilizer unit dampe n s m o v e m e n t o f t h e s e n s i n g a c t u a t o r. Pneumatic actuator and butterfly valve operation is the same as previously described. To reduce the amount of air being bled from the engine as altitude increases, an aneroid expands to vent more air pressure. The aneroid contracts as altitude decreases, venting less air pressure.

Flow Control Servo The flow control servo (Figure 21B-11), located in the aft fuselage aft of the rear pressure dome, meters deice system air to the signal port in the PRV. A DC signal from the bleed flow control positions a torque motor to reduce the signal pressure.

• A servoactuator that receives the output of the flow control servo The electropneumatic servo torque motor (Figure 21B-11) located adjacent to the pressure regulator, meters deicing system pressure to the pressure-sensing control. A DC electrical signal from the bleed flow control positions a torque motor to vent servo pressure overboard. If a high flow rate is selected by the BLEED control knob, the torque motor is positioned to vent a minimum amount of pressure from the open side of the pneumatic actuator. The remaining pressure overcomes the spring to drive the butterfly toward the open position. If a low flow rate is desired, the torque motor is positioned to vent a small amount of servo pressure overboard, the remaining pressure acting on the servoactuator to vent more pressure from the open side of the actuator. Spring force overcoming the servo pressure drives the butterfly toward the closed position. With the flow rate set, air venting is controlled by pressure sensed downstream of the valve. Pressure increase acts on a sensing actuator to vent more air; decreasing pressure results in

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21B-19

21B-20

NO. 1 HIGH PRESSURE SWITCH NO

NC

HIGH PRESSURE BLEED SHUT-OFF VALVE

NO. 2 HIGH PRESSURE SWITCH NO

NACELLE SHUT-OFF VALVE (OPEN)

OVERPRESSURE SWITCH NC

NC

NO CONTACT CLOSES P >75 PSI

NO B A OPEN WHEN PRESSURE > 65PSI

P19

A B C

OVERTEMPERATURE SWITCH

P7

A

B

B

P3

A

P13

B

C

CLOSES AT T > 550 ±10°

P13

A

B

P5

OPENS WHEN PRESSURE > 55 PSI

A B

C A3

D

A2 A1

SEE SHEET 2

B3 B2

NC A3

B1 D3 X2

D1

3

{

X1 X2 6121-K16

HBOV 0VERIDE BLEED SYSTEM 1 S3

K1 HIGH PRESSURE BLEED CONTROL RELAY (NORMALLY CLOSED)

1 2 3

OFF

4 5 6

BLEED ON BLEED SYSTEM CONT 1 5A (B5) 28V DC LEFT MAIN BUS CB1

CR 11

C1

2

4

A1 B3 CR6

B2

CR1

C3

C2

2

A2

NO

X1

D2

CR5

B1 NO 5 SEC TIME DELAY ON OPERATE B3 (AIR)

CR2 X1 X2 K3 TIME DELAY ON RELEASE A3 A2 A1 C1 X1 X1

B2

CR 19

B1 (GND)

X1 X2

A3 A2 A1

D1

X1

R3 X2

D3 3010-K2

E

K6

5 AUXILIARY LANDING RELAY

3261-K4 AIRFRAME AUTOMATIC SWITCH 1 (OFF) (SLOW) 2

2C1

CR20 (FAST) 3

3010-S3

AIRFRAME DEICE 28V DC (E8) 7.5 R. SEC CB4 BUS

4

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Figure 21B-12. Bleed-Air System Electrical Schematic—Series 200 (Sheet 1 of 2)

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CR9 NC

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SYSTEM OPERATION

NOTES

Selecting the BLEED switch to BLEED (Figure 21B-12) deenergizes the nacelle shutoff valve and both pressure regulator valves open and provides 28 VDC to the solenoid of time delay relay K3 through the normally closed contacts of the high-pressure bleed control relay (MSM Chapter 21). Power is also applied to time delay relay 3010K2 through the normally closed contacts of time delay relay K3 to control terminal C1 of deice delay relay 3010-K2. With both high-pressure switches closed, deice time delay relay 3010-K2 energizes and its normally closed contacts open to disconnect the ground from the deice system low-pressure warning switch. This prevents the DE-ICE PRESS caution light from illuminating during high-pressure shutoff valve operation. Five seconds from BLEED selection, time delay relay K3 energizes, and the high-pressure shutoff valves are energized open. With the valves open and the engines operating, high-pressure bleed air flows to the air-conditioning pack through the choking venturi, nacelle shutoff valves, and pressure regulator valve. Five seconds from the opening of the normally closed contacts of time delay relay K3, power is disconnected from deice delay relay 3010-K2 terminal C1. The relay is deenergized, reconnecting the ground to the deice system low-pressure warning switch.

Revision 3

21B-21

21B-22

NO. 1 HIGH PRESSURE SWITCH NO

OPEN WHEN PRESSURE > 65PSI

NC

C B A

HIGH PRESSURE BLEED SHUT-OFF VALVE

NO. 2 HIGH PRESSURE SWITCH NO

P18

NACELLE SHUT-OFF VALVE

OVERPRESSURE SWITCH

NC

A B C

OVERTEMPERATURE SWITCH

NO CONTACT CLOSES AT P >75 PSI A

P8

B

B

P9

A

P4

C B A

P12

CLOSES AT T > 550 ±10° A

B

P6

OPENS WHEN PRESSURE > 55 PSI

'B' 'A' 'E'

A3 A2 A1

3312-P3 'C'

12

LOGIC CIRCUIT

NO.1 BLEED HOT

B2

B3 B1

D3 X1

D2

C2

C3 X1 X2

3

6121-K15

3312-P31 7

LOGIC CIRCUIT

NO.2 BLEED HOT

MASTER CAUTION PANEL

28V DC CAUTION LIGHTS POWER

HBDV OVERRIDE OFF

BLEED SYSTEM CONT 2 (R5) 5A 28V DC RIGHT MAIN BUS CB2

CR12

2

X2

D1 C3

K2 HIGH PRESSURE BLEED CONTROL RELAY

1 2 3 4 5 6

BLEED ON 2

BLEED SYSTEM 2 S4

CR10

RIGHT BLEED NOTES: 1. IDENT CODE IS 2121, UNLESS OTHERWISE SPECIFIED. 2 FOR CONTINUATION, SEE CHAPTER 73, ENGINE ELECTRONIC CONTROL SYSTEM—ENGINE 1 AND 2. 3 FOR CONTINUATION, SEE CHAPTER 61, PROPELLER AUTO FEATHER CONTROL SYSTEM. 4 FOR CONTINUATION, SEE CHAPTER 30, AIRFRAME DE-ICE AND INDICATION. 5 FOR CONTINUATION, SEE CHAPTER 32, PROXIMITY SWITCH SYSTEM.

Figure 21B-12. Bleed-Air System Electrical Schematic—Series 200 (Sheet 2 of 2)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

When either nacelle upstream high-pressure bleed air reaches 65 psi on the ground or 55 psi in the air, the associated high-pressure switch opens and disconnects the ground from time delay relay K3 and deice time delay relay 3010-K2. Relay K3 deenergizes both highpressure shutoff valves closed and reconnects power to relay 3010-K2 terminal C1, energizing the relay and disconnecting the ground from the deice low-pressure warning switch. Low-pressure bleed air is then supplied to the system.

NOTES

Airborne with the AIRFRAME AUTO switch in SLOW or FAST, the bleed-air system remains in the high-pressure mode, controlled by the No. 1 high-pressure switch (65 psig). If an overpressure condition occurs, the wing bleed overpressure switch closes at 75 psi to energize the nacelle shutoff valve closed, isolating the affected bleed-air system.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RECIRCULATION FAN SWITCH ACTIVATES RECIRCULATION SYSTEM FAN WHEN SELECTED TO RECIRC.

AIR CONDITIONING OFF

CABIN DUCT TEMPERATURE INDICATOR INDICATES AIR TEMPERATURE IN CABIN DISTRIBUTION SUPPLY DUCT.

60 40 CABIN

20

0

MAN–AUTO SWITCH (2) AT MAN, TEMPERATURE CONTROL OF CABIN OR FLIGHT COMPARTMENT IS BY RELATED COOL–WARM SWITCH. AT AUTO, CONTROL IS DIRECTED TO ASSOCIATED AUTOMATIC TEMPERATURE CONTROLLER.

OFF

RECIRC

1

BLEED

80

DUCT 100 °C

MIN MAN

COOL

WARM

NOTE: F/A POSITION AT MAX COOL ON THE CABIN RHEOSTAT MOD 8/0807

MAX BLEED

COOL

WARM

AUTO

F/A CABIN

CABIN COOL–WARM SWITCH WITH RELATED MAN–AUTO SWITCH AT MAN, DIRECTLY CONTROLS PACK TEMPERATURE CONTROL VALVES WHEN HELD AT SPRING-LOADED COOL OR WARM POSITION.

2

TEMP CONTROL

FLT COMP

FLIGHT COMPARTMENT COOL–WARM SWITCH WITH RELATED MAN–AUTO SWITCH AT MAN. DIRECTLY CONTROLS TEMPERATURE TRIM VALVES WHEN HELD AT SPRING-LOADED COOL OR WARM POSITIONS. AUTOMATIC TEMPERATURE SELECTOR (2) ROTARY KNOB, WHEN POSITIONED AS DESIRED, INPUTS TEMPERATURE REQUIREMENT INTO AUTOMATIC TEMPERATURE CONTROLLER, WHICH MANIPULATES PACK TEMPERATURE VALVES OR TEMPERATURE TRIM VALVES TO MAINTAIN THE SELECTED LEVEL (MAN–AUTO SWITCH MUST BE AT AUTO.)

Figure 21B-13. Air-Conditioning System Controls

21B-24

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR-CONDITIONING SYSTEM GENERAL The air-conditioning system provides an inflow of temperature-controlled air to the cabin and flight compartment for heating, cooling, and ventilation. The system consists of a bleed-air system for each engine, an air-conditioning pack, and an air recirculation system. Conditioned air is supplied to the cabin and flight compartment through outlet grilles in the cabin ceiling and/or the lower sidewall vents flight compartment sidewalls. With the recirculation system operating, air is supplied to individually controlled outlets in the cabin, flight compartment, and lavatory. Bleed air from the engines is hot enough to meet all cold day requirements. On hot days, the bleed air is routed through an air cycle machine, where it is cooled to below ambient, using the pressure and energy of the air to drive the cooling turbine. When the air is cooled below its dew point, condensate is extracted by a condenser.

The cabin system ducts the air up to and along the cabin ceiling (or via the underfloor ducting depending on cabin air supply switch position), where it is discharged through lowvelocity air outlets along each side of the center aisle. The flight compartment system ducts air forward under the cabin floor to flight compartment side window demist outlets, directionally adjustable side console outlets, and fixed outlets at the rudder pedal foot wells. The air recirculation system improves cabin airflow and reduces the flow demand required of the air-conditioning system. Air is drawn from the flight compartment, through the avionics compartment, and into the underfloor duct by an electric recirculation fan in the duct, which forces the air aft to the mixer. Branch ducts just downstream of the fan direct part of the recirculation flow to a series of high-velocity air outlets (gaspers) located above each passenger position, in the lavatory, and on the pilot’s and copilot’s side panels. Air-conditioning system controls are shown in Figure 21B-13.

NOTES

COOLING SYSTEM General The major component of the cooling system is an air-conditioning pack located in the aft fuselage. It is supplied with bleed air from both engines and cools this air as necessary before distribution to the cabin and flight compartment. When heating is required, bleed air bypasses the refrigeration section of the pack. If cooling is required, the bypass valve is closed to divert bleed air through the refrigeration section. Temperature of the air supplied by the pack is controlled either automatically or manually from the flight compartment or the flight attendant’s panel if the cabin system is selected to AUTO.

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21B-25

21B-26

AIR CONDITIONING OFF

OFF

RECIRC

1

60 40 CABIN

20

0

BLEED

80

DUCT 100 °C

MIN MAN

COOL

2

WARM

MAX BLEED

COOL

WARM

AUTO

CABIN ZONE SENSOR INPUT

FLT COMPT DUCT HOT "FULL COLD" OVERRIDE SIGNAL

CABIN TEMP BIAS CONTROL

TEMP CONTROL

F/A CABIN

FLT COMP

FLIGHT COMPARTMENT TEMPERATURE CONTROLLER

CABIN TEMPERATURE CONTROLLER

FLIGHT COMPARTMENT ZONE SENSOR INPUT

"FULL COLD" OVERRIDE SIGNAL TO RAM-AIR SCOOP

DUCT GAUGE TEMP SENSOR DUCT TEMP SENSOR

APU

START BYPASS

BL AIR

DEICE SYSTEM

OPEN

BAFFLE BOX

TRIM VALVE ACTUATOR L ENGINE TEMPERATURE BLEED AIR TRIM VALVES BLEED-AIR SUPPLY DUCT

PACK TEMPERATURE CONTROL VALVES

ACM SUPPLY DUCT

APU BLEED-AIR VALVE

CONTROL VALVE ACTUATOR

R ENGINE BLEED AIR

RAM AIR OVERBOARD

CABIN DUCT

LEGEND

DUCT OVERTEMP SWITCH

PRESSURE REGULATING VALVE

FLT COMPT DUCT DUCT TEMP SENSOR

RECIRCULATED CABIN AIR

COMPRESSOR DISCHARGE OVERTEMPERATURE SWITCH

BLEED AIR HEAT EXCHANGER AIR

CONDENSER

COLD AIR

MIXING BOX WATER FILTER

HEAT EXCHANGERS

WATER NOZZLE

RECIRCULATED AIR CABIN CONDITIONED SUPPLY

WATER TRAP

FLIGHT COMPARTMENT CONDITIONED SUPPLY

FAN

AMBIENT OR RAM AIR

COMPRESSOR

WATER AIR CYCLE MACHINE (ACM)

Revision 3

AIR COND PACK HOT

EXPANSION TURBINE

ACM BYPASS DUCT

RAM-AIR DUCT

SHUTDOWN SIGNAL TO BLEED-AIR SYSTEMS

Figure 21B-14. Air-Conditioning Pack Schematic

ELECTRICAL POWER

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CAUTION LIGHT

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN DUCT HOT

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Air-Conditioning Pack The air-conditioning pack (Figure 21B-14), located in a bay in the aft fuselage, consists of an air cycle machine (ACM), primary and secondary heat exchangers, an exhaust duct, and associated piping and control valves. The pack is supplied with hot bleed air ducted in along the dorsal fin. According to temperature requirements, this air is cooled or passed as hot air to the mixer box.

Conditioned air leaving the mixer/condenser is separated by a diverter baffle for entry into the cabin and flight compartment air distribution systems.

NOTES

Air to be cooled is ducted through the primary heat exchanger, reducing its temperature, before entering the compressor side of the ACM where its pressure is increased. The compressor discharge is cooled through a pair of secondary heat exchangers and passes through the condenser where excess water is extracted. The air then enters the ACM expansion turbine, where it is cooled to low temperature and reduced in pressure for discharge into the mixer. The heat exchangers are cooled by ram air, ducted into the aft fuselage by dorsal fin ram inlets (supplemented by the vent door to relieve negative pressure when the airplane is on the ground), and drawn into a ram-air duct within the air-conditioning bay by an ACMdriven fan. The duct directs the air through the primary and secondary heat exchangers before discharging it overboard through the tail cone. Water piped from the condenser is sprayed into the ram-air duct to improve heat exchanger efficiency. Cold air discharged into the mixer is blended with cabin air forced into the mixer by the cabin recirculation system and hot bleed air injected upstream of the mixer via an ACM bypass duct. By varying the blended ratios of cold air to hot bypassed air, the desired conditioned air temperature is achieved. Pack temperature control valves, one positioned in the ACM supply duct and one in the ACM bypass duct, regulate the blending ratios in response to commands from the cabin temperature controller.

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21B-27

21B-28 CABIN VALVE

LINKAGE FLIGHT COMP VALVE

OPEN 90° 65° OPEN CLOSED 0°

MOUNTING PLATE

0° CLOSED

CCW STOP

CW STOP ACTUATOR ROTATION

FUNCTIONAL SCHEMATIC

FLIGHT COMPARTMENT CABIN BUTTERFLY VALVE BUTTERFLY VALVE

Figure 21B-15. Temperature Trim Valves

ACTUATOR

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

POSITION MICROSWITCH RANGE

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION AND OPERATION

NOTES

Temperature Trim Valve The temperature trim valve (Figure 21B-15) consists of two one-inch butterfly valves driven by an electrically operated actuator through mechanical linkage. All three components are installed on a mounting plate attached to the condenser/mixer in the aft fuselage, aft of the rear pressure dome. The valve regulates the temperature of the air supplied to the flight compartment by opening and closing the butterfly valves to control the flow of hot bleed air to achieve the desired temperature. The valves are controlled by an electrically operated actuator which modulates in response to automatic signals from the flight compartment temperature controller, or from direct electrical inputs from the manually operated switch. The mechanical linkage connecting the two butterfly valves to the actuator is designed to provide a sequence of valve openings and closings, in response to actuator rotation, to provide a complete temperature range for the flight compartment. If a warmer flight compartment temperature is desired, the flight compartment butterfly valve is opened to allow hot bleed air directly into the distribution system to achieve the temperature. The cabin butterfly remains closed. If a cooler flight compartment temperature is desired, the cabin butterfly valve is opened to allow hot bleed air directly into the cabin distribution system. The flight compartment butterfly valve remains closed. Since the cabin has not called for this extra heat, the pack temperature control valve is repositioned to allow less bypass. The resulting cooler pack discharge temperature restores cabin requirements (pack discharge plus added heat) and gives the flight compartment its cooler temperature.

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21B-29

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COLD FULL-TRAVEL MICROSWITCH

2.5-INCH THROTTLE VALVE

PACK BYPASS BUTTERFLY VALVE

OPEN

VALVE POSITION

1.5-INCH BYPASS VALVE

CLOSED

CLOSED

CCW STOP

THROTTLE BUTTERFLY VALVE ACTUATOR

MOUNTING PLATE

CW STOP ACTUATOR STROKE FUNCTIONAL SCHEMATIC

LINKAGE

Figure 21B-16. Pack Temperature Control Valves

21B-30

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Pack Temperature Control Valves

NOTES

The electrically actuated temperature control valve consists of a 1.5-inch diameter pack bypass butterfly valve operated in opposition to a 2.5-inch diameter throttle butterfly valve (Figure 21B-16). The two valves are mechanically linked to the actuator, and all three components are installed on a mounting plate attached to the heat exchanger in the rear fuselage aft of the rear pressure dome. The pack bypass valve is located in a bypass duct from the bleed-air duct; the throttle valve is in the outlet duct from the primary heat exchanger. The valve regulates the temperature of the air discharged from the pack by opening and closing of the pack bypass and throttle valves. The electrically operated actuator modulates in response to automatic signals from the cabin temperature controller or direct electrical inputs from the manually operated switch. The mechanical linkage connecting the two valves to the actuator provides a sequence of valve opening and closing, in response to rotation of the actuator, providing a complete temperature range from maximum cooling to maximum heating. For maximum cooling, the pack bypass valve is closed and the throttle valve is open. This allows the total flow from the engine bleedair system to pass through the dual heat exchanger and air cycle machine. When less cooling is required, the actuator moves the throttle valve toward the closed position and the bypass valve toward open to allow a proportional amount of the airflow to bypass the refrigeration circuit. At maximum heat, the throttle valve is fully closed, and the pack bypass valve is fully open.

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21B-31

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RAM-AIR OUTLET HEADER

VENT DOOR

RAM-AIR OUTLET

BLEED BYPASS TO MIXING BOX INLET DUCT (FROM ENGINE BLEED OR APU [SOO 8062]) PACK TEMPERATURE CONTROL VALVE MOUNTING BRACKET

PRIMARY

SECONDARY

OUTLET DUCT (TO ACM COMPRESSOR) RAM-AIR INLET INLET DUCT (FROM ACM COMPRESSOR) OUTLET DUCT (TO ACM TURBINE)

TRANSITION DUCT RAM-AIR INLET HEADER

WATER NOZZLE

Figure 21B-17. Heat Exchanger

21B-32

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Heat Exchanger

NOTES

General The heat exchanger (Figure 21B-17), located in the aft fuselage, is an all-welded aluminum alloy structure consisting of a housing with a finned core divided into primary and secondary units. Inlet and outlet headers on each side of the housing separate primary and secondary airflow. A circular transition duct, welded to a ram-air inlet header, functions as the ram-air inlet for the exchanger core. This duct is connected directly to the fan outlet end of the ACM. The primary exchanger inlet duct is connected to the bleed-air supply, and the outlet duct to the ACM compressor inlet. The secondary exchanger inlet duct is connected to the ACM compressor outlet and the outlet duct of the ACM turbine inlet through the condenser. Inward opening vent doors in the ram-air outlet header provide a means of pressure relief when the pressure differential in the air-conditioning bay reaches 0.4 psi. A spray nozzle in the ram-air inlet is connected to a water drain in the condenser and to another drain downstream of the condenser. Air tapped from the supply at the drains forces collected water to the nozzle to be sprayed into the heat exchanger inlet.

Heat Exchanger Operation The purpose of the heat exchanger is to lower the bleed-air temperature for the air cycle machine. Hot air from the bleed-air system flows to the primary heat exchanger. The airflow is partially cooled by heat transfer to ram air through the exchanger core and is then directed to the ACM compressor inlet. From the compressor, the air flows through the secondary heat exchanger where it is further cooled. The airflow is then directed through the condenser to and through the ACM turbine.

Revision 3

21B-33

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONDENSER MIXER

OUTLET TO CABIN AND FLIGHT COMPARTMENT

INLET FROM ACM TURBINE WATER DRAIN

INLET FROM RECIRCULATING FAN

CONDENSER OUTLET (TO ACM TURBINE)

WATER DRAIN

CONDENSER INLET (FROM HEAT EXCHANGER)

CONDENSER OUTLET

Figure 21B-18. Condenser/Mixer

21B-34

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Condenser/Mixer

NOTES

General The condenser/mixer (Figure 21B-18) is an allwelded aluminum alloy unit consisting of a finned core and inlet and outlet headers. It is bolted to structure in the aft fuselage. One end of the unit is the “mixer,’’ where cool air from the ACM, previously injected hot bypass bleed air, and recirculated cabin air mix to achieve the temperature required in the cabin. The other end of the unit is the condenser, where moisture is extracted from the air. Two drains collect extracted moisture and route it back to the heat exchanger inlet through a filter or the overboard drain.

Condenser/Mixer Operation Relatively cool conditioned air destined for the cabin passes through the finned core of the condenser. Warm air from the secondary heat exchanger passes over the cool core, where moisture in the air collects on the core. When droplets collect in water drains, they are routed back to the heat exchanger ram-air inlet or overboard drain.

Revision 3

21B-35

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TURBINE OUTLET (TO CONDENSER/ MIXER) FAN INLET SCREEN

HEAT EXCHANGER CONNECTING FLANGE

TURBINE HOUSING

COMPRESSOR INLET (FROM PRIMARY HEAT EXCHANGER)

COMPRESSOR HOUSING

FAN OUTLET DUCT

FAN HOUSING

OIL LEVEL SIGHT GAGE COMPRESSOR OUTLET OIL FILL (TO SECONDARY PLUG HEAT EXCHANGER)

Figure 21B-19. Air Cycle Machine

21B-36

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Air Cycle Machine General The air cycle machine (ACM) (Figure 21B-19) consists of a three-wheel assembly and separate housings for compressor, turbine, and fan rotors. The three housings are secured together at circular bolted flanges. The three-wheel assembly consists of a common shaft mounting a centrifugal compressor and turbine rotor on one end and an axial flow fan at the opposite end. The shaft rotates in a single bearing cartridge. The shaft seals are located to the front and rear of the bearing races to prevent oil leakage into the airflow. The bearings and the shaft are lubricated by oil from a sump in the turbine housing. Two wicks are radially preloaded against the shaft. The sump has drain and filler plugs and a sight gage. The dome-shaped compressor housing is flange-mounted on the turbine housing and contains a diffuser to direct the airflow leaving the centrifugal compressor. The compressor inlet is at the center of the domed housing and incorporates offset radial vanes to impart a swirling motion to the incoming air from the primary heat exchanger, thus correcting the impingement angle on the compressor rotor.

vents the ingress of foreign objects, thus protecting the fan rotor. The fan rotates in a steel shroud, which is an integral part of the inlet. The outlet section is cone-shaped with a tubular bore and air-straightening vanes that provide a divergent transition duct for ram-air delivery to the heat exchangers. The diffuser outlet mates with the ram-air inlet of the heat exchanger.

ACM Operation Cooled air from the primary heat exchanger enters the compressor inlet of the ACM, where it is compressed and then delivered at a higher pressure and temperature to the secondary heat exchanger through the condenser to the turbine inlet. Expansion of the air across the turbine rotor reduces the pressure, with a corresponding drop in temperature. Air leaving the turbine outlet is routed to the condenser/mixer. The turbine extracts energy from the airflow as it reduces the pressure to just above cabin pressure. Most of this energy is fed back to aid in driving the compressor. The remainder of the turbine energy drives the fan to ensure airflow through the heat exchanger when the airplane is stationary.

NOTES

The turbine housing is the main structural member of the ACM, supporting the compressor and fan housings on the end flanges and the bearings and three-wheel assembly in the bore. A turbine nozzle in the housing aids in the expansion of air leaving the turbine rotor, lowering the air temperature. Tube adapters on the housing connect the turbine inlet to the secondary heat exchanger and the turbine outlet to the condenser/mixer. The fan housing consists of two sections, one section serving as the air inlet and the other section providing the air outlet. The inlet section provides a circular, reverse-flow duct which directs ram air through the axial-flow fan rotor. A wraparound wire mesh screen pre-

Revision 3

21B-37

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR CYCLE MACHINE

Figure 21B-20. Compressor Discharge Overtemperature Switch

Figure 21B-21. Temperature Controllers

21B-38

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Compressor Discharge Overtemperature Switch

NOTES

The compressor discharge overtemperature switch (Figure 21B-20) is installed in the delivery duct from the ACM compressor. It consists of a normally open single-pole thermal switch with a bimetallic element. The switch closes when the compressor discharge temperature exceeds 405 ±10° F (207 ±5° C).

Overheat Condition If the temperature of the air discharged from the ACM compressor exceeds 207° C (450° F), the compressor discharge overtemperature switch closes to energize both HP bleed control relays closed with the following results: • The AIR COND PACK HOT caution light comes on. • Both HP bleed valves are deenergized closed. • Both nacelle shutoff valves are energized closed. When the overheat condition clears, the switch opens to restore system operation.

Temperature Controller The temperature controller (Figure 21B-21) for each system is located in the electrical equipment bay forward of the No. 1 relay panel. They consist of aluminum boxes housing electronic analog devices with printed circuit boards. A circuit in each controller is connected to its associated temperature control selector and duct and zone temperature sensors. Another circuit connects the applicable valve actuator and manual temperature control switch through relays. By comparing the input signals from the duct and zone sensors with the selection made, the controller completes a circuit to operate the applicable valve actuator. The actuator modulates the valves in response to the controller signals to supply conditioned air at the desired temperature.

Revision 3

21B-39

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DUCT OVERTEMPERATURE SWITCH

DUCT TEMPERATURE SENSOR

TEMPERATURESENSING BULB

PT FLT COM

CABIN

Figure 21B-22. Temperature Sensors and Switches

21B-40

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Duct Temperature Sensors The duct temperature sensors (Figure 21B-22) are negative coefficient thermistors in which the resistance varies inversely with the temperature of the air flowing through the duct. Each sensor provides a signal to the associated controller for automatic temperature control. The cabin duct temperature sensor is located in the cabin supply duct below the baggage compartment floor. The flight compartment duct sensor is in the flight compartment supply duct beneath the baggage compartment floor.

Zone Temperature Sensors The zone temperature sensors are similar in construction and operation to the duct sensors. They provide signals for automatic temperature control to the associated controller. The cabin zone temperature sensor with integral fan is located in the recirculation extraction dado, left side. The flight compartment zone sensor is located on the pilots consol, between the rudder pedals.

Temperature-Sensing Bulb The temperature-sensing bulb is an electrical resistance unit that transmits resistance variations proportional to temperature changes to the cabin DUCT temperature indicator (Figure 21B-23). The indicator reading allows for more precise temperature control when in the MAN mode.

resume air-conditioning system operation when the duct air temperature falls to 180° F (82° C).

OPERATION Air to be cooled is temperature-reduced through the primary heat exchanger and is then compressed through the ACM to boost the temperature and pressure. The compressed air passes through the secondary heat exchanger to further reduce its temperature, passes through the condenser where moisture is extracted, and then enters the ACM turbine. The energy extracted by the turbine drops the air temperature well below ambient and reduces the pressure to just above cabin ambient. At this point, bypassed bleed air is injected to achieve the desired temperature. The air then enters the mixing box, where it is blended with recirculated cabin air. Temperature of the air supply is regulated by opening and closing the pack temperature control valves to allow part of the hot bleed air to bypass the ACM and mix with the cool air discharged from the ACM. The valves are operated through mechanical linkage from a single actuator which positions the valves in response to signals from an automatic temperature controller (which can be monitored on the flight attendant’s panel) or from direct electrical inputs from a manually operated switch in the flight compartment.

NOTES

Duct Overtemperature Switches Each duct overtemperature switch consists of a normally open, single-pole thermal switch with a bimetallic element. When duct temperature exceeds 190° F (88° C), the switch closes to illuminate the appropriate cabin or flight compartment DUCT HOT caution light. Actuation of either switch also energizes a duct overheat control relay to switch the input to the temperature control valve actuator to the manual COOL command. The switch opens to

Revision 3

21B-41

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SOLENOID ACTUATOR HOUSING

ELECTRICAL CONNECTOR ACTUATOR COVER

ACCESS TO FILTER

FLOW CONTROL VALVE

SOLENOID SHOWN ENERGIZED FROM PRESSURE SWITCH

A B C

SOLENOID SHOWN DEENE FROM PRESSURE SWITCH

A B C

Figure 21B-23. Temperature Control System

21B-42

Revision 3

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TEMPERATURE CONTROL SYSTEM

NOTES

GENERAL Air-conditioning system temperature is controlled by two electrical subsystems, one for the cabin and the other for the flight compartment. The cabin subsystem controls temperature generated within the pack for the entire airplane; the flight compartment subsystem allows for changes in temperature between the two areas. Each subsystem is controlled by an independent set of switches on the AIR CONDITIONING panel, which can be positioned for manual or automatic operation. The cabin temperature control system utilizes a temperature controller that compares input signals from two sensors with the temperature selected and sends a signal to the pack temperature control valve actuator to modulate the control valves to achieve the desired temperature. The flight compartment temperature control system operates the same as the cabin system except the signal from the temperature controller is sent to the trim valve actuator to modulate the trim valves for the desired flight compartment temperature. Air-conditioning system controls on the overhead panel include a cabin DUCT temperature indicator and switches for temperature control (Figure 21B-23). Temperature in the cabin and the flight compartment can be controlled either automatically or manually with two switches labeled “MAN’’ and “AUTO.’’ The two manual temperature controls are momentary-on spring-loaded to center switches sharing a COOL–WARM legend with the automatic control CABIN and FLT COMP rotary variable resistors. The TEMP CONTROL MAN–AUTO switches are wired into the electrical circuits of the appropriate temperature controller.

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21B-43

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

UPPER DUCT

LOWER DUCT CABIN AIR SUPPLY

Figure 21B-24. Cabin Temperature Bias Control

21B-44

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

OPERATION Automatic Temperature Control Cabin In the automatic mode, the cabin MAN–AUTO switch (Figure 21B-23) is positioned to AUTO, and the CABIN automatic temperature control is rotated toward COOL or WARM, as required. Temperature of the cabin air is sensed by the cabin air duct temperature sensor in the inlet duct. Temperature of the air leaving the cabin is sensed by the cabin zone temperature sensor at the center left side of the cabin. The cabin zone sensor has an automatically powered fan to draw ambient cabin air over the sensor for a more accurate reading. These sensors are wired to form part of an electrical bridge circuit within the cabin temperature controller. An electrical input from the CABIN automatic temperature control selector, through the CABIN TEMPERATURE bias control, provides a variable in another leg of the bridge circuit. The controller compares the two sensed temperatures with the selected temperature, and the difference, in the form of electrical pulses, is transmitted to the pack temperature control valve actuator in the pack. The actuator responds to the signal by positioning the pack temperature control valve to regulate the flow of air at the desired temperature. Thus, the command output from the cabin temperature controller to the temperature control valve actuator brings the temperature at the zone sensor in the cabin in line with the selected temperature. The duct temperature sensor limits the air temperature to 35 to 160° F (2 to 71° C).

Flight Compartment Flight compartment air is initially at the same temperature as the cabin but can be adjusted by setting the FLT COMP automatic temperature control selector as required. This selection, plus the temperature sensed by the flight compartment duct temperature sensor in the aft section of the airplane below the baggage compartment floor and the flight compartment zone temperature sensor behind the pilot’s seat, forms a bridge circuit in the flight compartment temperature controller. The controller compares the two sensed temperatures with the selected temperature, and the difference, in the form of electrical impulses, is transmitted to the temperature trim valve actuator in the supply duct. If a warmer temperature than that in the cabin is desired in the flight compartment, the actuator positions a trim valve to meter hot air into the flight compartment supply duct. If a cooler temperature is desired, the flight compartment trim valve closes, and the cabin trim valve opens to route hot air into the cabin supply duct. If additional heat has not been selected, the cabin system causes the pack temperature control valve to reposition to restore cabin temperature, with the added hot air, and to provide the desired temperature in the flight compartment.

NOTES

Cabin Temperature Bias Control The CABIN TEMPERATURE bias control (Figure 21B-24), is located at the flight attendant’s forward station. When the flight compartment “cabin” rheostat is selected to the “F/A” position, the flight attendant gains full control of the cabin temperature.

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21B-45

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Manual Temperature Control Cabin In the manual mode, the cabin MAN–AUTO switch is positioned at MAN, and the required temperature is obtained by adjustment of the CABIN manual temperature control selector between COOL and WARM. In this mode, there are no limits in the supply temperature range. MAN selection also energizes a relay to initiate the following: • Isolate the hot and cold command signals from the cabin temperature controller to the pack temperature control valve actuator. • Complete the cool and warm circuits between the CABIN manual temperature control selector and the pack temperature control valve actuator.

• With the circuits complete, signals are sent to the actuator to position the valves as determined by the selection made at the FLT COMP selector. If a warmer flight compartment temperature is desired, the trim valves are positioned to add heat to the flight compartment supply duct. If cooler temperature is desired, the trim valves are positioned to add heat to the cabin supply duct. The cabin will readjust to the added heat in the AUTO mode. If the CABIN manual selection is made, the trim valves lock in the center position if the trim valve attempts to add heat to the cabin system. To release the lock, the cabin must be selected to AUTO, removing the ground from the locking relay.

NOTES

• With the circuits complete, signals are sent to the actuator to position the control valves as determined by the selection made at the CABIN manual temperature selector.

Flight Compartment In the manual mode, the MAN–AUTO switch is positioned at MAN, and the required temperature is obtained by adjusting the FLT COMP manual temperature control selector between COOL and WARM. Selection of MAN also energizes a relay to initiate the following: • Isolate the hot and cold command signals from the flight compartment temperature controller to the trim valve actuator in the supply duct. • Complete the cool and warm circuits between the FLT COMP manual temperature control selector and the trim valve actuator.

21B-46

Revision 3

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

STAGE THREE

Duct Overheat If the cabin or flight compartment supply duct air temperature exceeds 190° F (88° C), a duct overtemperature switch closes with the following results:

1. Pack valve goes full cold. 2. Microswitch causes return to center relay, driving trim valve to locked center position.

•

A circuit is completed to the master caution control box to illuminate a DUCT HOT caution light.

When the duct temperature falls to 180° F (82° C), the overtemperature switch opens to resume air-conditioning system operation.

•

The duct overheat relay K3 or K4 (MSM Chapter 21) is energized with the following results:

NOTES

•

Relay K1 or K2 is energized to isolate the automatic hot or cold commands from either temperature controller to the associated valve actuator.

•

A circuit is completed to provide a cool signal to the associated valve actuator, causing the actuator to reposition its valves to operate the pack in the fully cold mode.

Three stages of flight compartment overheat operation are as follows: STAGE ONE 1. Master CAUTION light illuminates. 2. Trim valve is disconnected for automatic and manual control. STAGE TWO Cabin in auto

Cabin in MANUAL

1. Trim valve adds heat to cabin ducts center.

1. Trim valve locks at center position.

2. Auto box senses heat and drives pack towards cold.

Revision 3

21B-47

21B-48

Canada Ltd.

Revision 3

Figure 21B-25. Conditioned Air Distribution

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AVIONICS COMPARTMENT COOLING

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CONDITIONED AIR DISTRIBUTION SERIES 200

NOTES

GENERAL Conditioned air from the pack is ducted into the fuselage at the center of the rear pressure dome bulkhead (Figure 21B-25). A longitudinal baffle in the duct directs 70% of the air to the cabin and 30% to the flight compartment via two ducts. The flight compartment duct is routed under the baggage compartment and cabin, where it splits into two ducts to feed the pilot’s and copilot’s subsystems. The cabin supply duct is routed under the baggage compartment, where it splits into an upper and lower duct for each side of the fuselage. The upper of these ducts supplies air to the cabin dado panel grilles, while the lower duct supplies, via sidewall risers, the upper cabin air outlets, as well as the cabin, lavatory, and flight attendant’s station gaspers. The flight compartment duct supplies air to sidewall grilles and gaspers in the flight compartment. Flight compartment air is also used for side window demisting. On airplanes with SOO 8069, conditioned air can be supplied to the cabin and flight compartment through an eight-inch universal ground air service connector located on the right side of the aft fuselage. The air supply from the ground connection is routed through a short flexible duct to join the cabin supply air duct in the aft fuselage. The air is then routed, via the condenser/mixer, to the cabin and flight compartment. A butterfly-type check valve is line-mounted in the ground conditioned air duct to prevent recirculated or airconditioning pack air from spilling overboard through the ground air connector.

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21B-49

21B-50

UPPER GRILLE GALLERY

GASPER GALLERY

DADO PANEL

RISER DUCT OVERTEMPERATURE SWITCH

MAIN CABIN SUPPLY DUCT

Figure 21B-26. Cabin Air Distribution Schematic

DUCT TEMPERATURE SENSOR

DUCT TEMPERATURE SENSING BULB

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT ATTENDANT CONTROLLED SHUTOFF VALVE

RESTRICTOR

TO FLIGHT ATTENDANT AND LAVATORY GASPERS

REAR PRESSURE BULKHEAD

Canada Ltd.

Revision 3

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN AIR DISTRIBUTION The supply of conditioned air for the cabin, consisting of approximately 70% of the air discharged from the air-conditioning pack, enters the cabin at the center of the rear pressure dome (Figure 21B-26). The air is routed below the cabin to dado panels above the cabin floor and, via sidewall risers, to the upper air outlets, which are located outboard of the passenger service (PSU) units. The supply duct splits at a wye duct to supply each side of the cabin. At the wye split, a secondary upper duct branches out from the lower duct and parallels it on each side of the fuselage. The upper duct supplies air to the cabin dado panels, while the lower duct supplies air to the upper air outlets and gaspers via sidewall risers. Incorporated into the upper duct is a motorized butterfly shutoff valve controlled by a two-position switch on the flight attendant’s panel. The switch positions are UPPER DUCT and LOWER DUCT. In the UPPER DUCT position, the shutoff valve shuts off air to the cabin dado panel, routes all air to the upper air o u t l e t s , a n d g ive s m a x i m u m f l ow t o t h e gaspers. When LOWER DUCT is selected, the shutoff valve opens to allow half of the air to the floor dado panels.

and an overtemperature switch, which are part of the temperature control system. A cabin duct temperature sensing bulb, located in the cabin supply duct downstream of the overtemperature switch, provides signals to a cabin duct tem-perature indicator on the AIR CONDITIONING panel in the flight compartment.

NOTES

In addition to normal cabin air conditioning, manually controllable air outlets (gaspers) are provided at each seat location, in the center ceiling above the forward flight attendant’s station, the lavatory, and the flight compartment. All gaspers, except those in the flight compartment, receive conditioned air from the cabin air supply duct via takeoff ducts from the same sidewall risers as the upper cabin (PSU) outlets. Approximately 60% of the cabin air is exhausted through slots located between the side and dado panels of the mid and aft cabin, passing under the cabin floor to be exhausted through the normal outflow valve. The other 40% of cabin air is exhausted through the air recirculation system. The delivery duct from the air-conditioning pack incorporates a duct temperature sensor

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21B-51

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FORWARD PRESSURE BULKHEAD

REAR PRESSURE BULKHEAD

SIDE WINDOW DEMIST OUTLETS

FLOW ADJUST LEVERS CABIN

TO LARGE FORWARD SIDE CONSOLE GASPERS

FIXED OUTLETS IN FOOT WELL

SIDE CONSOLE ADJUSTABLE OUTLETS

DUCT OVERTEMPERATURE SWITCH DUCT TEMPERATURE SENSOR

TO GASPERS BELOW WINDSCREEN CORNER PILLARS

CABIN SUPPLY

FLOW ADJUST LEVERS

ZONE TEMPERATURE SENSOR SIDE WINDOW DEMIST OUTLETS

Figure 21B-27. Flight Compartment Air Distribution

LAVATORY GASPER

FLIGHT ATTENDANT STATION GASPER

GASPER GALLERY

REAR PRESSURE BULKHEAD

FLIGHT COMPARTMENT

CABIN RESTRICTOR RISERS

LOWER CABIN AIR FLIGHT COMPARTMENT AIR LAVATORY

Figure 21B-28. Gasper System Schematic

21B-52

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Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT AIR DISTRIBUTION The supply of conditioned air for the flight compartment, consisting of approximately 30% of the air discharged from the air-conditioning pack, enters the cabin at the center of the rear pressure dome and is routed to the flight compartment via a duct under the cabin floor (Figure 21B-27). A wye duct splits the airflow into two individual but identical air distribution subsystems, one for the pilot and one for the copilot. Each of the subsystems consists of a side window demist twin outlet, a fixed outlet in the foot well, and an adjustable large outlet in the side console. Flow adjustment levers on the sill below the side window are cable connected to operate movable vanes in the supply ducts to permit adjustment of airflow. Both crewmembers are provided with one large and one small manually controllable (gasper) air outlet. The large gaspers are at the forward end of each side console. The small gaspers are located below the windshield pillars on both sides of the flight compartment.

that extend above the passenger service units (PSU) panels on each side of the cabin. From there the air is supplied to individual controllable outlets at each passenger seat. The risers also supply condi-tioned air to the cabin upper air outlet grilles. The upper cabin grilles require less airflow than the gaspers; therefore, a restrictor is installed in each riser between the gasper tapoff and upper grille gallery. An additional single riser supplies cabin duct air to a lavatory gasper and a flight attendant station ceiling gasper. The flight compartment gasper system is supplied with conditioned air from the flight compartment conditioned air supply duct. Two individual controllable outlets are provided for the pilot and the copilot. A large outlet is located at the forward end of each side console, and a smaller outlet is located below the windshield corner pillars.

NOTES

Air is exhausted from the flight compartment under the floor to the recirculation fan. From there the air is drawn back to the air-conditioning pack condenser/mixer by the recirculation fan. The delivery duct from the air-conditioning pack incorporates a duct temperature sensor and an overtemperature switch, which are part of the temperature control system.

GASPER AIR SYSTEM The gasper air system (Figure 21B-28) comprises two separate and independent systems, one for the cabin and one for the flight compartment. The one for the left system is supplied with conditioned air from the lower cabin conditioned air supply duct tapoffs from six sidewall risers (three on each side of the fuselage). The risers route air to the gasper galleries

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Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPT

AVIONICS COMPT

OUTFLOW VALVE CABIN

AIR EXTRACTION SLOTS

BAGGAGE COMPT CHECK VALVE

SLOT BETWEEN SIDE AND DADO PANEL

RECIRCULATION FAN REAR PRESSURE BULKHEAD

RECIRCULATION DUCT RECIRC FAN PWR 28 VDC LEFT SEC BUS

A

50A CR1

RECIRC FAN CONT 28 VDC 5A LEFT SEC BUS

OFF

2123K1 FAN CONTROL RELAY

B RECIRCULATION FAN

RECIRC

Figure 21B-29. Air Recirculation System Schematic

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIR RECIRCULATION SYSTEM The air recirculation system (Figure 21B-29) supplies air to the condenser/mixer of the airconditioning pack, where it is blended with conditioned air from the air cycle machine. The system consists of a recirculation fan, check valve, and associated ducting. The fan is controlled, through a fan control relay, by a switch marked OFF–RECIRC on the airconditioning panel. Electrical power is provided by the 28-VDC left secondary bus. The fan draws air back through the recirculation duct from under the flight compartment floor and from the cabin via adjoining air extraction ducts, located behind the top of the forward cabin dado panels, to the air-conditioning pack condenser/ mixer. The fan is located aft of the rear pressure bulkhead adjacent to the pack. The check valve is installed between the fan and the pack to prevent back-flow of conditioned air when the recir-culation system is not in operation.

NOTE The fan incorporates protective features for undervoltage, overcurrent, and overheat. The symptoms are those of a fan failure. The fan automatically shuts down when exposed to transient voltages of less than 22VDC for more than 0.10 second. The fan automatically restarts between 18 and 20 VDC. On occasion, an overcurrent condition will occur; the fan must be switched off for 20 seconds prior to restart. Overheat will also cause shut down, but it will automatically restart when the overheat condition clears.

Approximately 40% of the cabin air is recirculated; the other 60% is exhausted through slots located between the dado and side panels of the mid- to rear-cabin area. The exhausted air passes through the slots to the underfloor area and is exhausted through the normal outflow valve on the rear pressure bulkhead.

Operation Selecting the control switch to RECIRC completes a circuit to energize fan control relay 2123-K1. The relay then completes a circuit from the 28-VDC left secondary bus to the recirculation fan.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

LOW SPEED WARNING DEVICE EQUIPMENT THERMOSTATIC COOLING FAN SWITCH AIR EXTRACTION DUCT

FLT COMPT

AVIONICS COMPT

CABIN

FLT COMPT SENSOR

AVIONICS FAN

P1

LOW SPEED WARNING DEVICE FEABC

(20 SEC DELAY ON DIODE OUTPUT PIN F) WARNING LIGHT DS1

C B A

A3 A1

FAN

B3 B2 A2

A2

J2 P2

B1 A3 A1

X1 X2

K2

X1 X2

K1

S1 CLOSES AT > 35°C

COOLING FAN

10 28V DC L MAIN BUS P/O LEFT AVIONICS CIRCUIT BREAKER PANEL

NOTE: IDENT CODE IS 2126, UNLESS OTHERWISE INDICATED

Figure 21B-30. Avionics Compartment Cooling Fan

21B-56

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AVIONICS COMPARTMENT COOLING Avionics compartment cooling is accomplished by an avionics fan located under the cabin floor (Figure 21B-30). The fan draws air off each of the equipment shelves, via air circulation ducts, and discharges it under the cabin floor. The fan is a brushless, thermostatically controlled unit that is operational when DC power is applied to the airplane. In the event of fan failure, cooling by natural convection is adequate for interim operation. Fan failure is indicated by a light on the forward face of the flight attendant’s panel.

NOTE A 20-second delay is incorporated in the low-speed warning device on output pin F. This prevents a false fan failure indication while the fan is spooling up to operational speed after initial start.

OPERATION Normal The avionics cooling fan operates when the ambient temperature at the avionics rack exceeds 35° C (as detected by the thermostatic switch mounted on the upper avionics shelf). In this condition, switch S1 closes and relay K1 is energized. 28 VDC is applied through the fan circuit breaker and contacts of relay K1 and K2 to energize the fan and provide power to the low-speed warning device. Normal operating speed for the fan is 9,000 rpm.

Fan Failure Speed of the fan is monitored at pin C of the low-speed warning device by an output from the fan. If the speed of the fan drops to 6,000 rpm or less and remains at the reduced speed for more than 20 seconds, the low-speed warning device provides an output which energizes relay K2. Contacts of K2 disconnects power from the fan; K2 also maintains power to the low-speed warning device independent of relay K1. The low-speed warning device also turns on the diode warning lights DS1 to indicate fan failure.

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21B-57

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NOSE GEAR WELL FLIGHT COMPT

CARGO COMPT

CABIN

UNDERFLOOR AREA AFT PRESSURE DOME

LEGEND FORWARD PRESSURE BULKHEAD

PRESSURIZED

UNPRESSURIZED

Figure 21B-31. Pressurized Areas

3

2

6

CAB ALT

1 0 -1

4 5

1000 ft

7 8 9 10

BA RO in H G 31 30 29 28

ALT

BAR

CABIN ALTITUDE DUMP RATE F M A A U N L T AUTO CAB SET

NORM

Figure 21B-32. Pressurization Controls and Indicators

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PRESSURIZATION SYSTEM—SERIES 200

NOTES

GENERAL Pressurization of the airplane is dependent on three factors: (1) a positive, controlled airflow provided by the bleed-air system, (2) the flight compartment and cabin area being appropriately sealed, and (3) a controlled rate of air escaping from the fuselage. Air exhausted from the cabin and flight compartment is metered overboard through an outflow valve in the rear pressure dome. Valve operation depends on the setting of a pressure control unit on the overhead panel. A manually operated safety outflow valve in the forward pressure bulkhead can be used for a backup pressurization control or for smoke control. Pressurization is controlled by a normal outflow valve that modulates in response to electrical signals generated by a cabin pressure controller. The pressurized area of the fuselage is shown in Figure 21B-31. The system is monitored from the selector and indicator panels (Figure 21B-32). The selector panel provides for operation of the system in the automatic, manual, or dump modes. The indicator panel is used to monitor the system.

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21B-59

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PRESSURE RELIEF (SET TO RELIEVE AT 5.8 PSI)

CABIN AMBIENT

SIGNAL FROM PRESSURE CONTROLLER

CAB STATIC CAB

TORQUE MOTOR 18 PSI FROM DEICE SYSTEM

VENTURI

Figure 21B-33. Normal Outflow Valve—Series 200

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION AND OPERATION Normal Outflow Valve— Series 200 The normal outflow valve (Figure 21B-33) consists of a poppet and diaphragm assembly, a spring, an outer dome, a torque motor, and a differential pressure limiter. The valve is pneumatically operated and spring-loaded closed. Cabin pressure is applied to the inside of the poppet/diaphragm assembly and to the outer dome. Selecting the desired cabin altitude generates a DC signal to position the torque motor to meter suction produced by the venturi ejector to the outer dome. Thus, a pressure differential is created between each side of the poppet and diaphragm assembly. If the cabin altitude is less than selected altitude, the valve opens to bleed some of the cabin pressure, increasing the cabin altitude. A drop in cabin pressure is sensed as a reduction of poppet opening force, and the outflow valve moves toward the closed position to restore cabin pressure (decrease cabin altitude). If pressure differential between the cabin and ambient exceeds 5.8 ±0.15 psi, a differential pressure limiter in the outflow valve opens to connect the outer dome to ambient, bleeding pressure from the dome and opening the valve. The valve also provides negative pressure relief and opens if outside ambient pressure exceeds cabin pressure. The safety outflow valve on the forward pressure bulkhead is similar in operation but is controlled manually with a manual control needle valve on the control unit. Smoke can be removed from the flight compartment by opening the valve.

Revision 3

Venturi Ejector The venturi ejector is attached to the pressure dome adjacent to the outflow valve. Air pressure at approximately 18 psi from the airframe deicing system is directed to the venturi to produce suction, which is then applied to the outer dome of the normal outflow valve through the torque motor.

Safety Outflow Valve The safety outflow valve is similar in construction to the normal outflow valve. The safety outflow valve employs suction generated by venting one side of the valve to the slipstream; the suction is modulated by the manual metering valve. Thus, the system requires no electrical power or pneumatic pressure to function when the airplane is in flight.

Pressure Relief Valves Both the safety and normal outflow valves contain pressure relief valves which vent excess pressure if cabin pressure exceeds sensed ambient air pressure by 5.8 ±0.15 psi. The maximum regulated differential pressure permitted by the cabin pressure controller is 5.5 ±0.3 psi, which corresponds to an 8,000 foot cabin at 25,000 feet.

Negative Pressure Relief Valves Protection from negative pressure is also provided by the normal and safety outflow valves, which automatically open if the interior suction exceeds 0.1 psi.

NOTES

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN PRESSURE RATE-OF-CHANGE SELECTOR KNOB CONTROLS CABIN ALTITUDE RATE OF CLIMB/DESCENT WHEN SYSTEM IS IN SEMIAUTOMATIC MODE. FAULT LIGHT ILLUMINATES TO INDICATE SYSTEM MALFUNCTION. ALSO ILLUMINATES BRIEFLY DURING SYSTEM SELF-TEST.

CABIN ALTITUDE SETTING INDICATOR POINTER INDICATES CABIN ALTITUDE REQUIREMENT SET BY ALT KNOB.

MANUAL CONTROL KNOB DIRECTLY MODULATES SAFETY OUTFLOW VALVE SUCTION. TURNING CLOCKWISE OPENS OUTFLOW VALVE, CAUSING CABIN ALTITUDE TO INCREASE. CONTROL SENSITIVITY IS REDUCED AT LOWER ALTITUDES.

CABIN ALTITUDE KNOB INPUTS CABIN ALTITUDE REQUIREMENT INTO CABIN PRESSURE CONTROLLER.

3

2

6

CAB ALT

1 0 -1

4 5

1000 ft

7 8 9 10

BA RO in H G 31 30 29 28

ALT

BAR

CABIN ALTITUDE DUMP RATE F M A A U N L T AUTO CAB SET

INCR

NORM

BAROMETRIC CORRECTION INDICATOR

MODE SELECTOR SWITCH

BUG SHOWS CORRECTION SET ON BAR KNOB.

LEVER LOCKED SWITCH HAS THREE POSITIONS: AUTO — CABIN PRESSURE CONTROLLER IS ACTIVATED TO OPERATE FULLY AUTOMATICALLY OR SEMIAUTOMATICALLY, DEPENDING IN POSITION OF AUTOMATIC FUNCTION SWITCH.

BAROMETRIC CORRECTION KNOB INPUTS DESIRED BAROMETRIC CORRECTION INTO CABIN PRESSURE CONTROLLER AUTOMATIC FUNCTION SWITCH WITH MODE SELECTOR AT AUTO: NORM —SELECTS FULLY AUTOMATIC OPERATION CAB SET —SELECTS SEMIAUTOMATIC OPERATION

MAN —

CABIN PRESSURE CONTROLLER IS DEACTIVATED AND PRESSURIZATION CONTROL IS THROUGH SAFETY OUTFLOW VALVE VIA MANUAL CONTROL KNOB.

DUMP — NORMAL OUTFLOW VALVE IS HELD OPEN TO PREVENT AIRCRAFT FROM PRESSURIZING.

Figure 21B-34. Selector Panel

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Selector Panel The selector panel (Figure 21B-34) contains a combined cabin altitude and barometric indicator with ALT, BARO, and RATE knobs, an AU T O – M A N – D U M P s w i t c h , a C A B SET–NORM switch, a control needle valve marked “INCR,’’ and a FAULT light. The BARO knob is provided to set barometric pressure, the desired cabin cruising altitude/destination altitude is set with the ALT knob, and cabin altitude rate of change is set with the RATE knob. The RATE knob is rotated clockwise to increase the rate of change. When the ball on the knob is aligned with the index mark on the controller, the cabin rate of change is 500 fpm up and 300 fpm down. Movement full counterclockwise sets a minimum fpm and full clockwise sets the maximum rate of 0 to 2,500 fpm up and 1,500 fpm down.

The control needle valve is mechanically connected to the outer dome of the front safety outflow valve and to ambient. An arrow on the panel indicates that a clockwise selection opens the needle valve, venting the outer dome and increasing the cabin altitude. Full counterclockwise rotation closes the needle valve. The FAULT light provides indication of any fault in the pressurization control system. It illuminates for approximately two seconds when electrical power is first applied, indicating that dynamic self-testing is in progress. The light goes out if no fault is detected during the self-test, and passive continuous selftesting continues.

NOTES

The three-position AUTO–MAN–DUMP switch selects the mode of operation for pressurization control. In the AUTO position, control is completely automatic or semiautomatic, depending on the CAB SET–NORM switch position. In the MAN position, the desired cabin pressure is controlled by rotating the control needle valve toward INCR to increase cabin altitude or counterclockwise to decrease cabin altitude. In the DUMP position, the normal outflow valve is commanded to a fully open position and the airplane may be operated unpressurized. The CAB SET–NORM switch allows selection of pressurization in the CAB SET position or the cabin pressure controller to control pressurization automatically in the NORM position. In the CAB SET position, the crew selects the desired cabin cruising altitude and then the destination altitude (as descent begins) with the ALT knob. The cabin altitude rate of change can be varied with the RATE knob. In the NORM position, the destination altitude is selected prior to takeoff, and pressurization is automatic from takeoff to touchdown. In both cases, the BARO knob must be set to the correct barometric pressure.

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21B-63

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CABIN RATE OF CLIMB INDICATOR REGISTERS CABIN INTERIOR PRESSURE CHANGES WHICH ARE INDICATED AS A RATE OF CLIMB OR DESCENT

CABIN F OF

OFF

0

1

6

2

5 4

3

DIFF PSI

CABIN PRESS WARNING LIGHT ILLUMINATES TO INDICATE CABIN PRESSURE EXCEEDS 10,000 FEET.

30

0 2

20 14 12 10 ALT

4 6 8

FTx 1000

1

2

UP

0 DOWN

1

2

RATE

FPMx 1000

DIFFERENTIAL PRESSURE INDICATOR

CABIN ALTITUDE INDICATOR

REGISTERS DIFFERENCES IN INTERIOR AND EXTERIOR PRESSURE (NORMAL MAXIMUM 5.5 ± .3 PSI). POINTER RESTING AT OFF INDICATES NO POWER TO INDICATOR.

REGISTERS CABIN PRESSURE IN TERMS OF EQUIVALENT ALTITUDE. POINTER RESTING AT OFF INDICATES NO POWER TO INDICATOR.

Figure 21B-35. Indicator Panel Table 21B-1. PRESSURIZATION CONTROL SETTINGS SELECTOR PANEL SETTINGS AUTO/MAN/DUMP

CAB SET/NORM

RATE

ALT

BAR

MAN

AUTO

NORM

(INDEX MARK)

SET TO DESTINATION ALTITUDE

SET ATMOSPHERIC PRESSURE

FULL COUNTERCLOCKWISE

Electrical output from the computer controls suction applied to the normal outflow valve. Cabin altitude is maintained at, or as close as possible to, destination altitude without exceeding rate or differential pressure limits. AUTO

CAB SET

AS DESIRED

ENROUTE THEN DESTINATION

SET ATMOSPHERIC PRESSURE

FULL COUNTERCLOCKWISE

Electrical output from the computer controls suction applied to the normal outflow valve. Cabin altitude approaches and maintains enroute altitude. Destination altitude must be set at start of descent. MAN

N/A

N/A

N/A

N/A

MODULATE

Normal outflow valve is not used and is closed. Rotating manual dial meters low ambient pressure to modulate the safety outflow valve. Control effectiveness decreases as cabin pressure is reduced toward ambient. DUMP

N/A

N/A

N/A

N/A

N/A

Electrical output from the computer increases to apply full section to hold the normal outflow valve fully open.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Cabin Pressure Computer

NOTES

The cabin pressure computer, located on the avionics rack, controls the operation of the normal outflow valve. It compares dialed-in signals from the selector panel with cabin pressure, ambient pressure, and its computer schedule in relation to given information. It then sends an electrical signal to open or close the outflow valve to maintain the selected cabin pressure.

Indicator Panel The indicator panel (Figure 21B-35) consists of a DIFF pressure indicator, a cabin ALT indicator, and a cabin altitude RATE-of-change indicator. The indicators, receiving signals from the ADC, are used to monitor the pressurization control system in both the AUTO and MAN modes.

OPERATION General The pressurization control system operation is primarily electrical. The pressurized area of the fuselage is supplied with a relatively constant flow of conditioned air from the bleedair systems. Pressure in the fuselage is maintained by modulating the normal outflow valve to regulate the amount of air discharged to ambient. Both the normal and the safety outflow valves incorporate integral differential pressure limiters to open the valves when differential pressure exceeds a predetermined value. Both valves also open for negative pressure relief if the outside ambient pressure exceeds fuselage pressure. A summary of control settings for the various modes of operation is shown in Table 21B-1.

Revision 3

21B-65

21B-66

LEGEND B

DEICER PRESSURE CONTROL PRESSURE

AIR DATA COMPUTER

A CABIN PRESSURE CABIN AIR ATMOSPHERIC PRESSURE

NORM OPEN

ELECTRICAL

CABIN ALTITUDE DUMP

CABIN F OF

F OF 0

0 1

6

30 2

5 4 DIFF

3

3 2 1

2 4

20

6

14

12 10

ALT

4 5

FT 1000

8

2

1

UP

1

DOWN

1

RATE 7

CAB ALT

8 9

1000 FT

10

0

0

6

M A N

F A U L T

2

INCR

AUTO

RATE

ALT

CAB SET

BAR

NORM

HOUSING VENT

CAUTION LIGHT

18-PSI BLEED AIR FROM DEICING SYSTEM

DIGITAL COMPUTER

CABIN PRESS ALTITUDE PRESSURE SWITCH

TORQUE

CABIN AIR MOTOR

WEIGHT ON WHEELS POWER LEVERS ADVANCED INPUT CABIN AIR ATMOSPHERE

OUTFLOW VALVE

CABIN A

B

Figure 21B-36. Pressurization System Schematic

Canada Ltd.

Revision 3

P VALVE 5.8 PSI

VENTURI EJECTORINDUCED SUCTION

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

P VALVE 5.8 PSI

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

A schematic of the pressurization control system is shown in Figure 21B-36. A schematic of the electrical operation is shown in Figure 21B-39.

Automatic When electrical power is applied, the system self-tests, as indicated by illumination of the FAULT light on the selector panel, for approximately two seconds. If the system is faulty, the FAULT light remains illuminated.

is moved off its fully open stop and starts modulating to be ready to react instantly to the rapid changes, which occur following lift-off. When the landing gear relay is deenergized through the PSEU after lift-off, the valve modulates to maintain pressurization as governed by the computer program.

NOTES

If the self-test is satisfactory, the system is operational and can be controlled in either the CAB SET or NORM mode. In CAB SET, the crew selects the desired cabin cruising altitude, destination altitude (on descent), and cabin altitude rate of change, if desired, using the ALT and RATE knobs on the selector panel. In the NORM mode, only the destination altitude is set, using the ALT knob, and system operation is fully automatic with the information programmed into the computer. In both cases, the BARO knob must be set to correct atmospheric pressure.

On Ground When the airplane is on the ground with the weight on wheels and the power levers are retarded below 80% NH (12° above FLT IDLE), electrical power to pin R of the computer (MSM Chapter 21) is supplied through the normally open contacts of the energized landing gear relay 3261-K1 (the relay is energized through the PSEU). The system is now in the ground mode, and the normal outflow valve is fully open to prevent airplane pressurization.

During Takeoff When the power levers are advanced to 12° above FLT IDLE, power lever switch S1 closes to apply power to pin D of the computer. This, in turn, causes the computer to position the normal outflow valve to pressurize the airplane to 140 feet below ambient. The outflow valve

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

25,000

OPERATIONAL CEILING (25,000 FEET)

20,000

CABIN ALTITUDE—FEET

15,000

UNPRESSURIZED OPERATION

CABIN PRESS CAUTION LIGHT

10,000

MAXIMUM SCHEDULED CABIN ALTITUDE 8,000 FEET NORMAL PRESSURIZATION ENVELOPE 5,000 MP

/NOR

UTO

TE A XIMA

LE

EDU

E SC

UR RESS

MAXIMUM PRESSURE DIFFERENTIAL SCHEDULE (5.5 PSI)

R

APPO

SL

5,000

10,000

15,000

20,000

25,000

FLIGHT ALTITUDE—FEET

Figure 21B-37. Pressurization Envelope

Figure 21B-38. Forward Dump Manual Selector

21B-68

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

During Landing

Cabin Pressure Dump

The destination field altitude and atmospheric pressure are set with the ALT and BARO knobs on the selector panel. Pressurization is controlled automatically. If the selected field altitude is set higher than actual field altitude, the airplane lands unpressurized. If the field altitude is set below actual field altitude, the airplane lands pressurized. After landing, cabin altitude returns to field altitude for one minute at the rate selected before dumping cabin pressure to ambient.

Cabin pressure can be dumped by any of the following methods:

A schematic chart of the pressurization envelope is shown in Figure 21B-37.

Manual The pressurization level can be controlled with the safety outflow valve by selecting the AUTO–MAN–DUMP switch to MAN and setting the MAN control needle valve to achieve the desired pressurization. Rotating the control clockwise vents the pressure in the safety outflow valve to ambient through the needle valve. This opens the outflow valve, decreasing cabin pressure, which can then be regulated by rotating the MAN control to vary the amount of venting through the needle valve to obtain the desired pressurization level, as shown on the indicator panel. If the system is left at AUTO and is operating under automatic control, the resulting decrease in cabin pressure triggers the normal outflow to close when the MAN control knob is rotated to open the safety outflow valve. Thus, as long as the manual selection is for a higher cabin altitude (lower cabin pressure) than set on the automatic control, manual mode overrides the automatic selection.

Revision 3

• S e l e c t i n g t h e AU TO – M A N – D U M P switch to DUMP, causing the normal outflow valve to actuate to the fully open position • With CAB SET selected and the system in automatic mode, selecting the ALT knob to a cabin altitude above the flight altitude. This causes the cabin pressure to bleed off at a rate set by the RATE knob. • Selecting the forward dump selector to OPEN, causing the safety outflow valve to actuate fully open • Switching off bleed air

Forward Dump Manual Selector This selector (Figure 21B-38), located on the vertical plane of the copilot’s side console, provides for opening of the safety outflow valve to dump cabin pressure. The selector has positions labeled “NORMAL’’ and “OPEN.’’

Cabin Altitude Warning At a cabin altitude above 10,000 feet, or if a malfunction occurs in the system, a pressure switch on the avionics rack closes at a cabin pressure of 10 psi. This completes a circuit to illuminate the CABIN PRESS warning light.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

P4

P1 CHASSIS GND

C A B

S S

A

NORMAL OUTFLOW VALVE

CABIN SET HI K LO L AUTO J F P M N E G S T U

NORM MAN DUMP

}

S 28V DC L SEC BUS

V SELECTOR PANEL

VALVE-HI (10W) VALVE-LO (10W) +28V DC (25W) CABIN SET-NORMAL +28V DC CABIN DUMP TEST + 28V DC FAULT LAMP RETURN +10V REF POT RETURN BARO CORRECT SIGNAL CABIN RATE SIGNAL LANDING ALT SIGNAL CHASSIS GND PWR GND 28V DC—P.L. AT>12° AFI? 28VDC • WOW CABIN PRESSURE COMPUTER

P39 W e POWER C LEVER SWITCH NO NC S1

D3

J K A F C M N E G S T U V DC B D R

2

D2 D1

(28V DC OUTPUT—WOW) P/O 3261-K1

28V DC

R

3312P2 39

H

CABIN PRESS P/O MASTER CAUTION UNIT

PRESSURE P3 SWITCH C A B ACTUATES AT 10 PSIA (10,000)

A6

RELAY DRIVER

28V DC

LGWOW 1 NGWOW 1 RGWOW 1 LGWOW 2 NGWOW 2 RGWOW 2

LANDING GEAR SENSORS (WEIGHT ON WHEELS)

PROXIMITY SWITCH EELCTRONIC UNIT (PSEU) NOTE: 1. IDENT CODE IS 2131 UNLESS OTHERWISE INDICATED 2 SWITCH (SI) POSITION SHOWN WHEN POWER LEVER IS ADVANCED > 151/2° ABOVE FLT IDLE (SWITCH RELAXED)

Figure 21B-39. Pressurization Control—Electrical Schematic

21B-70

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CHAPTER 22 AVIONICS CONTENTS Page INTRODUCTION................................................................................................................. 22-1 GENERAL ............................................................................................................................ 22-3 PITOT-STATIC SYSTEM..................................................................................................... 22-5 General........................................................................................................................... 22-5 Pitot Heads..................................................................................................................... 22-7 Static Vent Plates ........................................................................................................... 22-7 Airspeed Indicator ......................................................................................................... 22-9 Inertial Vertical Speed Indicator .................................................................................... 22-9 Barometric Altimeter................................................................................................... 22-11 Altitude Preselector/Alerter......................................................................................... 22-11 AIR DATA SYSTEM.......................................................................................................... 22-13 General ........................................................................................................................ 22-13 Air Data Computers..................................................................................................... 22-14 Emergency Warnings................................................................................................... 22-15 Avionics Standard Communications Bus (ASCB) ...................................................... 22-15 ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS) ..................................... 22-17 General ........................................................................................................................ 22-17 Description and Operation........................................................................................... 22-17 Power Requirements.................................................................................................... 22-19 System Test.................................................................................................................. 22-21 ELECTROMECHANICAL FLIGHT INSTRUMENT SYSTEM...................................... 22-23

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Attitude Director Indicator (ADI) ............................................................................... 22-23 Radio Altimeter ........................................................................................................... 22-25 Horizontal Situation Indicator (HSI)........................................................................... 22-27 Radio Magnetic Indicator (RMI)................................................................................. 22-29 ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) ............................................. 22-31 General ........................................................................................................................ 22-31 EADI ........................................................................................................................... 22-35 EHSI ............................................................................................................................ 22-39 EFIS Controller ........................................................................................................... 22-41 Instrument Remote Controller..................................................................................... 22-41 System Test.................................................................................................................. 22-41 MISCELLANEOUS AND STANDBY FLIGHT INSTRUMENTS .................................. 22-43 Clock ........................................................................................................................... 22-43 Standby Magnetic Compass ........................................................................................ 22-45 Standby Altimeter........................................................................................................ 22-45 Standby Attitude Indicator .......................................................................................... 22-47 Turn-and-Slip Indicator ............................................................................................... 22-47 SPERRY DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM ................................. 22-49 General ........................................................................................................................ 22-49 Automatic Flight Control System (AFCS).................................................................. 22-51 NAVIGATION EQUIPMENT ............................................................................................ 22-61 General ........................................................................................................................ 22-61 Navigation System Receivers ...................................................................................... 22-61 COMMUNICATIONS EQUIPMENT ................................................................................ 22-65 Audio Integration ........................................................................................................ 22-65 Service Interphone....................................................................................................... 22-65

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VHF COMM Radios ................................................................................................... 22-67 VHF NAV Radios........................................................................................................ 22-69 Emergency Operation.................................................................................................. 22-69 Passenger Address System .......................................................................................... 22-71 Audio Integrating, Passenger Address, and Attendant’s Interphone—Series 300............................................................................................... 22-73 Static Discharge Wicks................................................................................................ 22-75 MISCELLANEOUS EQUIPMENT ................................................................................... 22-75 Transponder................................................................................................................. 22-75 Cockpit Voice Recorder............................................................................................... 22-77 Flight Data Recorder ................................................................................................... 22-79 Ground Proximity Warning System ............................................................................ 22-81 Emergency Locator Transmitter .................................................................................. 22-83 Weather Radar ............................................................................................................. 22-85 Flight Management System......................................................................................... 22-87 AFCS INTERFACE BOX (NON-EFIS) ............................................................................. 22-89 Description .................................................................................................................. 22-89 TRAFFIC ALERT AND COLLISION AVOIDANCE (TCAS) ......................................... 22-92 System Description...................................................................................................... 22-92 TRAFFIC ADVISORY DISPLAYS ................................................................................... 22-97 Flat-Panel VSI/TRA Display....................................................................................... 22-97 ELECTRICAL POWER SOURCES .................................................................................. 22-98 MAINTENANCE CONSIDERATIONS ............................................................................ 22-98 General......................................................................................................................... 22-98 Pitot-Static System ...................................................................................................... 22-98 Air Data System........................................................................................................... 22-98

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Automatic Flight Control System................................................................................ 22-98 Flight Instruments and Navigation System Components ............................................ 22-99 Communications.......................................................................................................... 22-99 FUNCTIONAL CHECKS................................................................................................... 22-99 FAULT ANALYSIS ............................................................................................................ 22-99 LIMITATIONS.................................................................................................................... 22-99

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ILLUSTRATIONS Figure

Title

Page

22-1

Avionics Component Locations ............................................................................. 22-2

22-2

Pitot-Static System................................................................................................. 22-4

22-3

Pitot Head............................................................................................................... 22-6

22-4

Static Vent Plate ..................................................................................................... 22-6

22-5

STATIC SOURCE Selector ................................................................................... 22-6

22-6

Airspeed Indicator.................................................................................................. 22-8

22-7

Inertial Vertical Speed Indicator ............................................................................ 22-8

22-8

Barometric Altimeter ........................................................................................... 22-10

22-9

Altitude-Alerting Profile...................................................................................... 22-10

22-10

Air Data System................................................................................................... 22-12

22-11

Attitude and Heading Reference System............................................................. 22-16

22-12

AH-600 Strapdown AHRU.................................................................................. 22-18

22-13

FX-220 Flux Valve .............................................................................................. 22-18

22-14

IMU Sensor Configuration .................................................................................. 22-20

22-15

AHRS Controller ................................................................................................. 22-20

22-16

Attitude Director Indicator .................................................................................. 22-22

22-17

Radio Altimeter Receiver-Transmitter................................................................. 22-24

22-18

Horizontal Situation Indicator ............................................................................. 22-26

22-19

Radio Magnetic Indicator .................................................................................... 22-28

22-20

Sperry EFIS Data Flow........................................................................................ 22-30

22-21

Sperry EFIS Displays (Typical)........................................................................... 22-32

22-22

EADI Symbology ................................................................................................ 22-34

22-23

EADI Caution and Failure Annunciations........................................................... 22-34

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22-24

EHSI Symbology ................................................................................................. 22-36

22-25

EADI/EHSI Composite Display .......................................................................... 22-38

22-26

EHSI Caution and Failure Annunciations ........................................................... 22-38

22-27

EFIS Controller.................................................................................................... 22-40

22-28

Symbol Generator Failure Indicator .................................................................... 22-40

22-29

Davtron Clock...................................................................................................... 22-42

22-30

Standby Magnetic Compass ................................................................................ 22-44

22-31

Standby Altimeter................................................................................................ 22-44

22-32

Standby Attitude Indicator................................................................................... 22-46

22-33

Turn-and-Slip Indicator ....................................................................................... 22-46

22-34

Sperry DFZ-8000—Electromechanical Flight Instruments................................. 22-48

22-35

Sperry DFZ-8000—Electronic Flight Instruments.............................................. 22-50

22-36

AFCS Controller.................................................................................................. 22-52

22-37

GA Buttons .......................................................................................................... 22-54

22-38

Control Wheel Switches ...................................................................................... 22-54

22-39

Advisory Display ................................................................................................. 22-58

22-40

Servoactuators...................................................................................................... 22-58

22-41

NAV Control Head............................................................................................... 22-60

22-42

Marker Beacon Lights ......................................................................................... 22-60

22-43

MARKER SENS Panel........................................................................................ 22-60

22-44

ADF Control Head............................................................................................... 22-62

22-45

Audio Integrating System .................................................................................... 22-64

22-46

Audio Control Panels........................................................................................... 22-66

22-47

VHF COMM Control Head ................................................................................. 22-68

22-48

Passenger Address Components and Controls .................................................... 22-70

22-49

Static Discharge Wicks ........................................................................................ 22-74

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22-50

XPDR Control Head ............................................................................................ 22-74

22-51

CVR Monitor Panel ............................................................................................. 22-76

22-52

Cockpit Voice Recorder System .......................................................................... 22-76

22-53

FDR/ELT Control Panel ...................................................................................... 22-78

22-54

Flight Data Recorder System............................................................................... 22-78

22-55

Flight Data Recorder Block Diagram .................................................................. 22-79

22-56

Ground Proximity Warning System..................................................................... 22-80

22-57

Emergency Locator Transmitter .......................................................................... 22-82

22-58

Weather Radar Indicator ...................................................................................... 22-84

22-59

Flight Management Computer (Typical) ............................................................. 22-86

22-60

Gables Mode C, Mode S Transponder and TCAS Control Panel........................ 22-94

22-61

Typical VSI/TRA Display ................................................................................... 22-96

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CHAPTER 22 AVIONICS

INTRODUCTION The Dash 8 avionics covered in this chapter is found in Chapters 22, 23, and 34 of the Maintenance Manual and includes the pitot-static system, Sperry digital integrated flight control system, flight instruments, navigation equipment, and communications equipment. The standard avionics package and the optional electronic flight instrument system are described; the user should consult applicable supplements in the AFM, the Maintenance Manual and vendor manuals for additional information and information on specific systems not included in this chapter.

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BUFFET UNIT

FLIGHT ATTENDANT'S SEAT AVIONICS COMPARTMENT WARDROBE

P/A AFCS INTERFACE

115-V, 400-Hz COOLING FAN CONTROLLER CABIN PRESSURE RAD ALT 1 RAD ALT 2 ADC 1 ADC 2

PSEU

FGC 1 FDAU

HF

WARDROBE

GPWS FGC 2

FLIGHT ATTENDANT'S SEAT

DME 1

ADF 1 ADF 2

DME 2

VHF/NAV 1 VHF/NAV 2 AFCS

AUDIO INT ATC 1 VHF COMM 1

ATC 2

VHF COMM 2

Figure 22-1. Avionics Component Locations

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GENERAL This chapter goes into detail on the operation of the avionics installed in the Dash 8. Avionics malfunctions may not be simple and easily diagnosed, and gold-plate or shotgun removeand-replace maintenance techniques can be very expensive, especially if the replacement is not really necessary. With rapidly changing and increasingly more sophisticated electronic avionics equipment installed in today’s airplanes, it is imperative that the line mechanic fully under stand the operation of all systems and the limits within which a pilot can operate those systems. This does not require that maintenance personnel be instrument pilots; instead, it requires them to be knowledgeable enough of instrument flight to understand inflight malfunctions as reported by the flight crew, functionally test and duplicate the malfunction if possible, and take the proper steps to correct the malfunction with minimum expense to the operator.

clocks. The altimeters are covered with the air data system, and the ADIs, HSIs, and EFIS are covered under the digital integrated flight control system. Navigation equipment includes the flight management system and the marker-beacon, VOR, DME, and ADF receivers. Other systems covered are the transponder, cockpit voice recorder, flight data recorder, ground proximity warning system, emergency locator transmitter, and weather radar. The Dash 8 communications included in this chapter are the VHF COMM, service interphone, and passenger address systems. Figure 22-1 shows the locations of the avionics components.

The following avionics systems are covered in this training manual. The pitot-static system includes two pitot heads, two static vent plates with three static ports per plate, and a dual air data system. The Sperry digital integrated flight control system includes the air data system, the auto matic flight control system (AFCS), the attitude and heading reference system (AHRS), and either the standard electromechanical ADIs and HSIs or the electronic flight instrument system (EFIS). The flight director function can be used independently of the autopilot with the pilot steering the airplane to satisfy the flight director commands as programmed, or the autopilot may be coupled to automatically steer the airplane to satisfy the flight director commands. The yaw damper system operates independently of the autopilot and may be engaged with or without the autopilot engaged. Dash 8 flight instruments covered include the standby instruments, airspeed indicators, vertical speed indicators, inclinometers, and

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MANIFOLD

ASI

PILOT'S PITOT HEAD

IVSI

PILOT'S INSTRUMENT PANEL STATIC

MANIFOLD

STBY ALT ENGINE INSTRUMENT PANEL

PILOT'S INSTRUMENT PANEL STATIC

SOURCE

ALTERNATE

NORMAL

AIR DATA COMPUTER NO. 1

COPILOT'S PITOT HEAD

ASI

PILOT'S STATIC SOURCE SELECTOR (NORMAL SELECTED)

IVSI

SOURCE

NORMAL

AIR DATA COMPUTER NO. 2

COPILOT'S STATIC SOURCE SELECTOR (NORMAL SELECTED) PILOT'S STATIC

LEFT STATIC PORT

RIGHT STATIC PORT

ALTERNATE STATIC COPILOT'S STATIC RIGHT STATIC HEAD

LEFT STATIC HEAD

CABIN ALTITUDE CONTROLLER

LEGEND PILOT'S PITOT COPILOT'S PITOT PILOT'S STATIC COPILOT'S STATIC ALTERNATE STATIC

Figure 22-2. Pitot Static System (Sheet 1 of 2)

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NOTES

PITOT-STATIC SYSTEM GENERAL The pitot-static system serves the computers of a Sperry ADZ-810 air data system, a standby altimeter, airspeed indicators, inertial vertical speed indicators, and the cabin pressurization system. Provision is made for addition of optional equipment. Two pitot heads provide impact air pressure, and two static vent plates provide static air pressure. The pitot heads and the static vent plates are electrically heated to prevent icing. Consult the Maintenance Manual for information on the moisture trap test points located at the lowest points in the lines. The pitot-static system is shown in Figure 22-2 (Sheets 1 of 2 and 2 of 2).

FWD

PILOT'S STATIC PORT

DRAIN HOLE

DRAIN HOLES (TOP AND BOTTOM)

ALTERNATE COPIlOT'S STATIC STATIC PORT PORT

PITOT PORT

RIGHT SIDE STATIC HEAD

PITOT HEAD (LEFT AND RIGHT SIDES SIMILAR)

PILOT'S STATIC PORT

PILOT'S STATIC PORT

FWD

COPILOT'S STATIC PORT

ALTERNATE STATIC PORT

LEFT SIDE STATIC HEAD

FWD COPILOT'S STATIC PORT

ALTERNATE STATIC PORT

LEFT SIDE (EARLY AIRPLANES)

Figure 22-2. Pitot-Static System (Sheet 2 of 2)

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Figure 22-3. Pitot Head

Figure 22-4. Static Vent Plate

22-6

Figure 22-5. STATIC SOURCE Selector

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PITOT HEADS

NOTES

One pitot head (Figure 22-3) is mounted on each side of the fuselage nose. The pitot heads provide independent supplies as follows: • Left pitot head—No. 1 air data computer and the pilot’s airspeed indicator • Right pitot head—No. 2 air data computer and the copilot’s airspeed indicator Both pitot heads are electrically heated by 28 VDC. Pitot heat is controlled by two PITOT S TAT I C s w i t c h e s o n t h e ove r h e a d I C E PROTECTION panel (see Chapter 30, “Ice and Rain Protection,’’ for additional information). Water drains are provided in the pitot heads.

STATIC VENT PLATES One static vent plate (Figure 22-4) is mounted on each side of the fuselage forward of the pitot head. There are three static ports in each static vent plate. Each port serves a separate static air system and is interconnected with the corresponding port of the static vent plate on the opposite side of the airplane. The three static vent systems are: • Pilot’s static air system • Copilot’s static air system • Alternate static air system Refer to Figure 22-2 to see the interconnection of the systems. The alternate static ports supply the cabin alt i t u d e c o m p u t e r. W h e n e i t h e r S TAT I C SOURCE selector (Figure 22-5) is positioned from NORMAL to ALTERNATE, the associated flight instruments for that side are supplied with alternate static air pressure and the normal system is cut off. Static vent plate anti-icing heat is supplied through the PITOT STATIC heat switches on the overheat ICE PROTECTION panel.

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AIRSPEED DRUM

400 AIRSPEED SCALE

0

50 20

WHITE INDEX BUG

30 100

300

AIRSPEED POINTER

MAXIMUM OPERATING SPEED (VMO) POINTER YELLOW INDEX BUG

150 200

WHITE INDEX BUG

YELLOW INDEX BUG CONTROL KNOB

Figure 22-6. Airspeed Indicator

Figure 22-7. Inertial Vertical Speed Indicator

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AIRSPEED INDICATOR

NOTES

The airspeed indicators (Figure 22-6) sense pitot air pressure to indicate airspeed and static pressure to indicate maximum operating speed (V MO ). Airspeed indications are shown using a white airspeed pointer against a 50- to 400knot scale and a rotating drum which reads from 0 to 100 knots. The maximum operating speed is preset and compensated for altitude changes. V MO indicated is 242 KIAS from sea level to 14,000 feet MSL and decreases linearly to 207 KIAS at 25,000 feet MSL for the 100 Series aircraft and 243 KIAS from sea level to 17,000 feet, decreasing linearly to 214 knots at 25,000 feet for the 300 Series aircraft. Index bugs are provided for pilot V-speed reference.

INERTIAL VERTICAL SPEED INDICATOR The inertial vertical speed indicators (Figure 22-7) sense the rate of change of static pressure supplied by the pitot-static system. Each instrument has a yellow pointer and range markings which indicate from 0 to 4,000 feetper-minute climb or descent. The pointers are mechanically driven and compensated by a dash-pot assembly to eliminate pointer lag. With TCAS installation, this is replaced by an electronic instrument.

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Figure 22-8. Barometric Altimeter

AUDIO ALERT (FOR 0.75 SEC)

ALERT LIGHT ON UNTIL RESET

+1000 FT ALERT LIGHT ON

PRESELECT ALTITUDE ERROR

DEPARTURE

CAPTURE

+250 FT

ALERT LIGHT OFF ALERT LIGHT OFF

0 FT –250 FT

SELECTED ALTITUDE

AUDIO ALERT

888 88

CAPTURE ALERT LIGHT ON

DEPARTURE

ALTITUDE FEET

–1000 FT ALERT LIGHT ON UNTIL RESET

Figure 22-9. Altitude-Alerting Profile

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BAROMETRIC ALTIMETER The barometric altimeter (Figure 22-8) provides a servoed counter drum and pointer display of barometrically corrected altitude. The indicator provides the following displays: • Altitude alert annunciator—An altitude alert light is located in the upper-right corner of the instrument bezel. It illuminates when the airplane is within 1,000 feet of the preselected altitude and extinguishes when the airplane is within 250 feet of the preselected altitude. The light comes on any time the airplane departs more than 250 feet from the selected altitude and remains on until reset or the airplane returns to within 250 feet of the selected altitude. • Failure warning flag—The flag is displayed when the error between the displayed altitude and the altitude signal received is too great, the ADC goes invalid, or the barometric altimeter power is lost (26 VAC, 400 Hz). • Altitude counter—The four-drum counter displays altitude from 0 to 60,000 feet. A negative (NEG) altitude shutter obscures the 10,000 and 1,000 digits of the counter to annunciate altitudes below sea level. The zero position on the ten-thousands drum is black-and-white crosshatch to alert the pilot to altitudes of less than 10,000 feet.

ALTITUDE PRESELECTOR/ALERTER The altitude preselector/alerter is used to preselect an altitude from 0 to 60,000 feet. Actual altitude is compared with the preselected altitude, and visual and/or aural (optional) alerts are generated by the ADC in two situations, depending on whether the system is in acquisition ( capture) mode or deviation (after-capture) mode. When in acquisition mode and a point 1,000 feet above or below the preselected altitude is reached, a one-halfsecond to one-second tone (2,900 Hz) is sounded, and the barometric altimeter alert light illuminates. When the airplane is within 250 feet of the preselected altitude, the alert light goes out, and the system switches to deviation mode. If ± 250 feet of the preselected altitude is exceeded, the alert light and tone are triggered. The deviation alert is the same as the acquisition alert. When the airplane returns to within 250 feet of selected altitude, the light extinguishes. If the airplane does not return to the selected altitude, the preselected altitude must be reset to extinguish the altitude alert light. Figure 22-9 illustrates the altitude- alerting profile. The Altitude Preselector displays dashed lines if the the active Air Data Computer fails.

• BARO set knob and barometric counters—The BARO set knob is used to set barometric pressure, which is displayed in inches of mercury and millibars. This correction is also sent to the ADC, which in turn updates the altimeter displays. • Pointer—The pointer displays altitudes between 1,000-foot levels on a scale with major indices every 100 feet and minor indices every 20 feet.

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22-12 NO. 1 TEMPERATURE PROBE (LOCATED LEFT WING)

NO. 1 NAV DATA

PITOT PRESSURE PILOT'S STATIC PRESSURE ALTIMETER

1013

2992

6

4

5

BARO

NM

SEL

KTS MIN

PWR

ALTITUDE ALERTER (ON ENGINE INSTRUMENT PANEL) ADC TEST 1

BANK

ALT

AP

NAV

CAT 2

VS

YD

APP

STBY

FLC

M TRIM

VNAV

CPL

B/C

ADVISORY DISPLAY

0

1

8 28 720 2 7

MB

1013

6 BARO

AIRSPEED EXCEEDS VMO

ASCB

ALT IN HG 3 2992

5

PSEU RUDDER* OVERSPEED

4

ALTITUDE ALERT ALTIMETER SETTING ALTITUDE ALT FAIL FLAG

AIRSPEED EXCEEDS VMO

NO. 2 ADC

AIRSPEED EXCEEDS 140 KTS AIRSPEED EXCEEDS 130 KTS STATIC PRESSURE AIRSPEED EXCEEDS 140 KTS ENCODED ALTITUDE

PITOT PRESSURE COPILOT'S STATIC PRESSURE ALTIMETER

OVERSPEED WARNING RUDDER * PSEU AIR CONDITIONING FLIGHT SPOILERS NO. 2 ATC TRANSPONDER

* THE RUDDER SIGNAL IS IN SERIES THROUGH BOTH ADCs. NO. 2 NAV DATA

TO OBTAIN REDUCED RUDDER HYDRAULIC PRESSURE, BOTH ADCs MUST BE SERVICEABLE. OPERATING SPEED FOR THESE SWITCHES IS 150 KTS IN THE 300 SERIES.

Figure 22-10. Air Data System

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NO. 2 TEMPERATURE PROBE (LOCATED LEFT WING)

NO. 1 & NO. 2 AHRS NO. 1 & NO. 2 EFIS

ADVISORY DISPLAY ALTITUDE SELECT ALTITUDE DISPLAY O'SPEED SIGNAL HSI SELECT

FGC CONTROLLER (ON GLARESHIELD PANEL) 9

AIRSPEED EXCEEDS 130 KTS AIRSPEED EXCEEDS 140 KTS

NO. 1 ATC TRANSPONDER FLIGHT SPOILERS AIR CONDITIONING

NO. 1 & NO. 2 FGC

TEST 2 (ON PILOT'S SIDE PANEL) HDG

STATIC PRESSURE NO. 1 ADC

ALTITUDE SELECT ALTITUDE DISPLAY O'SPEED SIGNAL HSI SELECT

HLD NAV PRE 12 MLS

888 8 8

AIRSPEED EXCEEDS 140 KTS

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

9 0 1 8 28 720 2 7 MB ALT IN HG 3

ALTITUDE ALERT ALTIMETER SETTING ALTITUDE ALT FAIL FLAG

ENCODED ALTITUDE

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AIR DATA SYSTEM

NOTES

GENERAL The ADZ-810 air data system (Figure 22-10) is a combination of sensors, instruments, and two computers that display parameters associated with the air mass (altitude, airspeed, vertical speed, and temperature). The system supplies this data to the flight control system, navigation system, warning system, barometric altimeter, transponder, flight recorder, and other airplane subsystems. System components include: • Two air data computers • Two barometric altimeters • Altitude preselector/alerter A four-line advisory display provides SAT/TAS and air data command information as well as sensor and system status messages and warnings. The advisory display is covered later in this chapter under Automatic Flight Control System.

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AIR DATA COMPUTERS

• Vertical speed

Two air data computers (ADC) independently perform computations and conversions based on altitude, airspeed, and temperature and provide the data to airplane systems. Self-monitoring circuits check that the input data is reasonable and that the computed output is correct. Information into the ADC is from analog and digital sources. Computations are digital, and output is in analog and digital form.

• Static source error correction • Digitized pressure altitude for mode C (transponder) • V MO • Overspeed warning During power-up, the following takes place:

ADC output is available to airplane subsystems, including the flight guidance computers, the attitude and heading reference system, the advisory displays, and, if installed, the EFIS. Such data is transmitted in a multiplexed, digital form via the avionics standard communications bus (ASCB), and, depending on the information, may be displayed on either an analog or digital instrument. The ASCB is an information bus that allows each of these systems to receive the information it needs from other systems and to share information that it produces with those systems. Both ADCs are normally operating when power is on their electrical buses. Information from only one is selected for flight guidance computer input and for the advisory display. Selection of the HSI SEL switch on the flight guidance controller panel selects the pilot’s or copilot’s HSI and ADC information for use by both flight guidance computers.

• Normally, No. 1 ADC is primary and is indicated by the left HSI edgelight. • If during power-up No. 1 ADC fails, a momentary caution is flashed on line one of the advisory display, indicating the failure. No. 2 ADC automatically takes over, and the right HSI edgelight is indicated on the flight guidance controller. • An ADC failure after power-up requires manual selection to be made to the serviceable ADC by one press of the HSI button on the flight guidance controller. The ADC has an operating range of –1,000 to 60,000 feet altitude and 30 to 450 knots.

NOTES

The ADC receives pneumatic information from the pitot-static system. Total air temperature and static air temperature are computed from inputs from the total air temperature probe. These data are then used to calculate all other air data parameters, including: • Indicated airspeed • True airspeed • Pressure altitude • Barometric correction • Barometrically corrected altitude

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EMERGENCY WARNING

NOTES

When the air data computers sense that the airplane speed is above the calculated V MO , a warning horn in the cockpit sounds an intermittent 1,000-Hz tone. Because the method used by the ADC to compute V MO is more accurate than that of the airspeed indicator, the result is that the warning horn may occur up to 6 knots indicated airspeed after the airspeed indicator has reached V MO. For Series 100 aircraft, V MO is 242 KIAS from sea level to 14,000 feet and decreases linearly to 207 KIAS at 25,000 feet. For Series 300 aircraft, V MO is 243 KIAS from sea level to 17,000 feet, down to 214 KIAS at 25,000 feet. The airplane overspeed warning circuits can be checked using the ADC test switch located on the pilot’s side panel.

AVIONICS STANDARD COMMUNICATIONS BUS (ASCB) The ASCB is an information bus that allows each avionics system to receive the information it needs from other systems and to share information that it produces with those systems. These two paths are called Bus A and Bus B, and each consists of two-wire pairs denoted as “data’’ and “clock.’’ Two bus controllers are used to manage all data transfer activity. With dual interconnections and dual bus controllers, bus availability from the essential to highly essential level is achieved. The bus controllers are located in the flight guidance computers. One bus controller only is active at any time. The other acts as a backup and assumes control of the bus when required due to failure of the active controller.

Revision 2

22-15

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLUX VALVE (LOCATED IN RIGHT WINGTIP)

FLUX VALVE (LOCATED IN LEFT WINGTIP) MAGNETIC HEADING REFERENCE PILOT'S AHRS CONTROLLER (LOCATED ON CENTER CONSOLE)

OFF SLOW

SLOW

FAST

DG

FAST

HDG

VG ERECT

DG

MAGNETIC HEADING REFERENCE

AUX PR OFF TEST

SLOW

FAST

SLOW

FAST

DG SLEW

DG

FAST

HDG

VG ERECT

DG

COPILOT'S AHRS CONTROLLER (LOCATED ON CENTER CONSOLE)

AUX PR TEST

FAST

DG SLEW SLAVE

BASIC

SLAVE

BASIC

MODE SELECTION AND HEADING REFERENCE MODE SELECTION AND HEADING REFERENCE

NO. 1 AHRU

NO. 2 AHRU

ASCB

EFIS ADIS AND HSIS

AIR DATA COMPUTERS

FLIGHT GUIDANCE COMPUTERS

ADVISORY DISPLAYS

(IF INSTALLED) ATTITUDE

PILOT'S STANDARD ADI AND HSI (IF INSTALLED)

3

24

6 9 12

24

0

27 30

33

6 9 12

15

18 21

PILOT'S RMI

(IF INSTALLED)

3

27 30

0

HEADING

HEADING AND ATTITUDE

COPILOT'S STANDARD ADI AND HSI

A D F

18 21

33

WEATHER RADAR

VOR

15

HEADING AND ATTITUDE

VOR

HEADING

A D F

COPILOT'S RMI

Figure 22-11. Attitude and Heading Reference System

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Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS) GENERAL The dual AHRS is the primary attitude and heading reference source. Several operational modes are provided for redundancy in the event of certain system failures. The standard system consists of dual attitude and heading reference units, dual AHRS controllers, and dual flux valves. Pitch, roll, and heading are provided to the standard flight instruments or EFIS and to the AFCS. Attitude and heading data are also supplied to other systems (e.g., weather radar antenna and RMIs).

only by momentarily pressing and releasing the HDG/DG pushbutton on the AHRS controller. The green DG then illuminates on the controller, indicating that the AHRS is in DG mode. AHRS operation in DG is similar in behavior to a free directional gyro, subject to drift and turn errors. For this reason, AHRS operation in the DG mode results in degraded heading accuracy.

NOTES

DESCRIPTION AND OPERATION General The standard AHRS operating modes are “normal’’ for attitude and “slaved’’ for heading (Figure 22-11). These modes are entered automatically after completion of initialization, provided that all system components and signals are valid. In normal mode, ADC true airspeed compensates for acceleration-induced error sources. The slaved mode receives magnetic heading reference from the flux valve. Two reduced-performance modes are also available: “BASIC’’ for attitude and DG (directional gyro) for heading. BASIC is annunciated in green on the AHRS controller and entered automatically when ADC true airspeed becomes invalid. AHRS operation in basic mode results in attitude system behavior similar to that of a conventional vertical gyro, which has pitch-and-roll erection cutoffs and is subject to drift and acceleration errors. For this reason, AHRS operation in basic mode results in degraded attitude accuracy. DG mode disables automatic slaving of the heading. Entry into this mode can be achieved

Revision 2

22-17

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-12. AH-600 Strapdown AHRU

(H) D L4

(C) E

28-VDC, 400-Hz EXCITATION

L1 F

SHIELD

L2 L3 B C

800-Hz SIGNAL OUTPUT

A

Figure 22-13. FX-220 Flux Valve

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Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Attitude and Heading Reference Unit (AHRU)

NOTES

The AHRU is an attitude and heading sensor (Figure 22-12) which provides digital and selected analog outputs for airplane guidance, control, and display. The AHRU is capable of 360° operation in the roll and heading axes and ± 90° in the pitch axis. The strapdown mechanization also provides additional outputs of airplane dynamic quantities such as body rates and accelerations.

FX-220 Flux Valve No. 1 flux valve (Figure 22-13) is in the left wingtip, and No. 2 is in the right wingtip. The flux valves detect the magnitude and direction of the earth’s magnetic field and convert them to electrical information which is used to align the directional gyros to magnetic north.

POWER REQUIREMENTS Power required to operate the system is 28 VDC. Each AHRS has two 28-VDC circuitbreakers, labeled No. 1 AHRS and No. 2 AHRS for primary power; and No. 1 AHRS AUX and No. 2 AHRS AUX for auxiliary power. If primary power falls below 18 VDC for 200 milliseconds or more, the system automatically switches to auxiliary power and the amber AUX PWR light on the controller illuminates.

Revision 2

22-19

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

X

P

Z ACC (FORCE REBALANCE)

Q

P

R

X

Y

Z P GYRO

Q GYRO

R GYRO

Y

Q X ACC (TOROID)

Y ACC (TOROID) Z

Figure 22-14. IMU Sensor Configuration

AMBER

AMBER

DG

FAST

AUX PR

HDG

VG ERECT

GREEN

OFF SLOW

SLOW

FAST

DG

TEST

FAST

DG SLEW SLAVE

RED

BASIC

GREEN

Figure 22-15. AHRS Controller

22-20

Revision 4

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Inertial Measurement Unit (IMU)

• SLAVE annunciator—The annunciator indicates a flux valve slaving failure.

The IMU measures the body axis angular rates and linear accelerations and converts them to DC signals. Three rate gyros and accelerometers (Figure 22-14) are contained in the IMU. Each rate gyro is aligned to a principal airplane axis. Two types of accelerometers are used: a servo unit in the Z axis and two toroidal units in the X and Y axes. The toroids are liquid-levels which are self-temperature-compensating. The servo unit is also temperature-compensated and requires no special drive signals.

• BASIC annunciator—The annunciator indicates when the AHRU is in the reversionary mode. The basic mode, which occurs automatically, is a reversionary mode in the attitude channel used to maintain performance in the event that true airspeed data is not available.

AHRS Controller A line drawing of the AHRS controller is shown in Figure 22-15. Following are the AHRS controls and their operation: • DG SLEW switch—The switch provides for selection of bidirectional dual rate heading slew commands. The switch is spring-loaded to the center OFF position and is active only in DG mode. • HDG/DG switch—The momentary pushbutton switch provides for the selection of the directional gyro mode or the heading-slaved mode. • DG annunciator—The annunciator indicates when the AHRU is in the directional gyro mode. It is a reversionary mode that is selected if a flux valve fails. • VG ERECT switch—The momentary pushbutton switch provides for the selection of the fast vertical erection mode. • FAST annunciator—The annunciator indicates when the AHRU is performing a fast vertical erection. • TEST switch—The momentary pushbutton switch provides for selection of the maintenance test mode for the AHRS. • AUX PR annunciator—The annunciator indicates when the AHRU is in the auxiliary power mode.

Revision 4

SYSTEM TEST The AHRS performs automatic self-test when power is first applied. The test lasts five seconds and provides the following visual indications on the ADI and HSI: • 10° pitchup • 20° right wing down • North heading, turning at 3° per second toward east • All AHRS controller annunciators on • ATT flag valid for 2.5 seconds, then invalid • HDG flag valid for 2.5 seconds, then invalid The flags remain invalid until initialization is complete (complete initialization requires 3 minutes). The test sequence may be initiated manually at any time (including in flight) by momentarily pressing the TEST pushbutton on the AHRS controller. The indications on the displays are the same as those seen during the automatic power-up test except that the pitch, roll, and heading test values are added to the existing airplane flight conditions. Upon completion of the five-second test, the system returns to the correct values, the flags clear, and the annunciators extinguish. Basic and DG modes are not affected by the test. The test sequence may be extended by holding t h e T E S T bu t t o n i n f o r l o n g e r t h a n five seconds. The test ends when the button is released, and the flags and annunciators clear if all data is valid.

22-21

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT DIRECTOR WARNING FLAG

ROLL SCALE

ROLL ATTITUDE POINTER DECISION HEIGHT ANNUNCIATOR

GO-AROUND ANNUNCIATOR GA

DH

PITCH-AND-ROLL COMMAND BARS FAST

FD

ATT

20

20

ATTITUDE WARNING FLAG

10

10

ATTITUDE SPHERE

10

10

GLIDE-SLOPE POINTER

SPEED COMMAND POINTER AIRCRAFT SYMBOL DECISION HEIGHT DISPLAY

ATTITUDE TEST SWITCH

20

SLOW

200 DH

20

1570

RAD ALT

DIM

RADIO ALTITUDE DISPLAY

DH SET/ TEST

INCLINOMETER

EXPANDED LOCALIZER POINTER

DECISION HEIGHT SET KNOB AND DIM CONTROL

Figure 22-16. Attitude Director Indicator

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Revision 4

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ELECTROMECHANICAL FLIGHT INSTRUMENT SYSTEM ATTITUDE DIRECTOR INDICATOR (ADI) General The electromechanical ADI is a Sperry AD550C (Figure 22-16). It combines the attitude sphere display with computed steering information to provide the commands required to intercept and maintain a desired flight path. It also contains an eyelid display, expanded localizer indication, glide-slope indication, digital radio altitude display, decision height set and display, and inclinometer. GA and DH annunciators are located in the top-left and topright corners, respectively, of the instrument.

Description and Operation Roll attitude is presented by a moving pointer against a fixed scale. The sphere roll angle is one-to-one: for every degree the airplane actually rolls, the sphere reflects it degree-fordegree. In contrast, the pitch presentation is nonlinear to allow a better resolution around the zero-degree pitch reference. The attitude test button (ATT) provides a self-test of the attitude presentation. The eyelid is a small ring around the sphere w h i c h a lwa y s p r ov i d e s p o s i t ive a t t i t u d e identification regardless of the sphere. It maintains proper ground-sky relationship and aids recovery from unusual attitudes. T h e a t t i t u d e f l a g p r o v i d e s A H R S - va l i d information, ADI input power status (26 VAC, 400 Hz), and a discrepancy monitor between the AHRS pitch/roll information and what the ADI is actually showing. The flight guidance computers drive the ADI crossbars. When no flight director mode is selected, the bars are biased out of view. An FD flag identifies failures within the FZ-800 computer.

Revision 4

Localizer deviation is amplified and presented as a reference for landing. The display is automatically reversed when the back course mode is selected on the flight director mode controller. The pointer is biased from view until a valid ILS frequency has been tuned. Radio altitude is shown in digital format. The readout displays altitude in five-foot increments between –20 and 200 feet and in ten-foot increments between 200 and 2,500 feet. Decision height is selected using a bezelmounted knob and is displayed in ten-foot increments to 990 feet. The altitude display blanks above 2,500 feet, and the DH display can be blanked by setting a DH below zero feet. A knob concentric to the DH set knob allows dimming of the RAD ALT and DH displays, as well as DISTANCE and COURSE displays on the HSI, and the ALTITUDE PRESELECT display. Pressing the DH set knob initiates a self-test of the radio altimeter system (radio altitude reads 100 feet).

Glide-slope deviation is presented by a pointer against the right vertical scale. The green vertical band on both sides of the index mark represents the Category II window. The GS pointer biases out of view upon loss of a glide-slope receiver valid. An inclinometer shows slip or skid. A go-around (GA) annunciator is located in the upper left corner of the bezel, and a DH annunciator is located in the upper-right corner. The green GA illuminates when the flight director is in go-around, and the amber DH illuminates when the airplane is below decision height. A relative speed indicator is displayed on the left side of each ADI. The speed indication is a “fast’’ or “slow’’ indication relative to 1.3 times the airplane stall speed and is intended for use in an approach or in slow flight. The speed control indicators are operated by the stall warning system. The fast-slow commands are displayed by a pointer against the left vertical scale. The pointer is biased out of view if invalid inputs are received or the display is not being used.

22-23

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ZERO ALTITUDE ADJUSTMENT

Figure 22-17. Radio Altimeter Receiver-Transmitter

22-24

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RADIO ALTIMETER

NOTES

The radio altimeter system provides absolute altitude from 2,500 to 0 feet, decision height selection, failure annunciation, and internal self-test. The display is blank if the radio altitude is above 2,500 feet AGL. If the radio altitude valid input is invalid, a series of dashes appear. The EADI and ADI display radio altitude in the lower-right corner. Figure 22-17 shows the radio altimeter receiver-transmitter. The receiver-transmitter provides a DC output voltage which is proportional to the airplane absolute altitude above the terrain. In addition, it provides radio altitude trip points, an indicator warning flag output, and an auxiliary radio altitude output. The precision output supplies absolute altitude information to the flight directors.

Revision 2

22-25

22-26

COURSE SELECTOR KNOB SETS THE COURSE POINTER AND COURSE COUNTER

SPOILERS ROLL OUTBD

FLIGHT

PULL UP GPWS TEST

COURSE

COURSE

PULL UP GPWS TEST

ROLL INBD

ENGINE FAIL

12:02

ENGINE PROPELLER GROUND RANGE

GROUND

ENGINE FAIL TAXI

BELOW G/S

53

BELOW G/S

12:02

ENGINE

53

HEADING

HEADING

ON

12

OFF

OFF

SETS THE HEADING BUG

COMPASS WARNING FLAG

3

3

W

N A V

V E R T

12

POINTER IS IN VIEW ONLY WHEN TUNED TO A LOCALIZER FREQUENCY AND DISPLAYS GLIDE-SLOPE DEVIATION. THE AIRCRAFT IS BELOW GLIDEPATH IF POINTER IS DISPLACED UPWARD. EACH DOT REPRESENTS 0.4 DEGREE DISPLACEMENT ABOVE OR BELOW GLIDE-SLOPE CENTER.

24

W

V E R T

HEADING BUG DIST

E

12

24

33

GLIDE-SLOPE POINTER E

N A V

PROVIDES DIGITAL DISPLAYS OF DME DISTANCE TO OR FROM A STATION IN NAUTICAL MILES.

6

30

33

PROVIDES MAGNETIC BEARING TO A SELECTED GROUND BASED NAVAID (VOR).

DISTANCE INDICATOR

21

HD

C

BEARING POINTER DIST

6

INDICATES NAVIGATION INPUTS FROM VOR, TACAN, OMEGA, OR LOCALIZER ARE NOT VALID OR POWER FAILURE.

00 0

INDICATES DEVIATION IN DEGREES FROM A SELECTED COURSE CENTERLINE DURING VOR 00 0 OR ILS COURSE OPERATION

C

INDICATES SELECTED COURSE. COURSE

FIXED HEADING INDEX FOR COMPASS CARD.

COURSE DEVIATION BAR

COURSE POINTER NAVIGATION WARNING FLAG

FORE LUBBER LINE

INDICATES POWER FAILURE OR DIRECTIONAL GYRO FAILURE.

HD

DISPLAYS MAGNETIC COURSE INFORMATION ON A DIAL WHICH ROTATES WITH THE AIRCRAFT THROUGHOUT 360 DEGREES.

30

ROTATING HEAD DIAL

15

COURSE COUNTER PROVIDES A DIGITAL READOUT OF A SELECTED COURSE IN DEGREES.

THIS NOTCHED ORANGE HEADING BUG IS POSITIONED ON THE ROTATING HEADING DIAL BY THE HEADING KNOB, AND DISPLAYS PRESELECTED COMPASS HEADING. IN THE HEADING MODE. THE ADI COMMAND BAR WILL DISPLAY THE PROPER BANK COMMANDS TO TURN TO AND MAINTAIN THIS SELECTED HEADING.

GLIDE-SLOPE WARNING FLAG INDICATES LOSS OF VALID GLIDE-SLOPE SIGNAL OR POWER FAILURE.

15

21

RECIPROCAL COURSE POINTER INDICATES RECIPROCAL OF COURSE SELECTED.

COURSE/TRACK DEVIATION SCALE

AFT LUBBER LINE

COMPASS SYNCHRONIZATION ANNUNCIATOR

TO-FROM POINTER

DURING VOR, TACAN, OR OMEGA OPERATION, EACH DOT REPRESENTS 5 DEGREES DEVIATION FROM CENTERLINE BY COURSE DEVIATION BAR, DURING ILS OPERATION, EACH DOT REPRESENTS 1/14 DEGREES FROM CENTERLINE BY COURSE DEVIATION BAR.

FIXED RECIPROCAL HEADING INDEX FOR COMPASS CARD

CONSISTS OF THE SYMBOL OR (DOT OR CROSS) DISPLAYED IN A WINDOW. WHEN THE COMPASS SYSTEM IS IN THE SLAVE MODE, THE DISPLAY WILL OSCILLATE BETWEEN AND , INDICATING THE HEADING DIAL IS SYNCHONIZED WITH GYRO STABLIZED MAGNETIC HEADING.

TWO FLAGS 180° APART INDICATE DIRECTION OF SELECTED COURSE.

•

•

+

+

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

HEADING SELECTOR KNOB

Canada Ltd.

Revision 4

Figure 22-18. Horizontal Situation Indicator

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

HORIZONTAL SITUATION INDICATOR (HSI)

indications from the glide-slope receiver causes the VERT flag to come into view over the glide-slope display.

General The Sperry RD-550A HSI combines numerous displays to provide a map-like presentation of airplane position. The indicator displays airplane displacement relative to VOR, localizer, and glide-slope beams and heading with respect to magnetic north. Figure 22-18 shows the RD-550A indications.

NOTES

Description and Operation The HSI has remote controls located on the glareshield. They are used to select heading and course. A conventional heading dial displays heading from the AHRS against a reference index (lubber line) at the top center. Additional indices are located at 45° on each side of the top reference. The heading flag comes into view for a failure of the AHRS or loss of HSI input power (26 VAC, 400 Hz), or from a discrepancy monitor between AHRS heading information and what the compass card is displaying. VOR bearing is displayed on an arrow-shaped pointer read against the heading dial. S t a n d a r d VO R - I L S r a d i o i n f o r m a t i o n i s displayed. Left-right deviation, VOR to-from indication, and a NAV flag are displayed. DME distance and selected course are displayed as digital readouts. These displays are dimmed using the concentric knob on the ADI DH set knob. Selected heading and selected course are set using the remote controllers in the glareshield. A slaving indicator for the AHRS heading output is located in the lower-left corner of the HSI. Glide-slope deviation is displayed by a pointer against a vertical scale. Loss of valid data or

Revision 4

22-27

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

LUBBER LINE

HEADING FLAG— INDICATES LOSS OF VALID SIGNAL OR POWER FAILURE

NO. 2 ADF/VOR POINTER

RMI COMPASS CARD

HD G 33

E

21

9

24

3

15

S

ADF

V O R

MODE INDICATOR— DIRECTION OF ARROW INDICATES MODE SELECTED

12

V O R

NO. 1 ADF/VOR POINTER

N

W

32

ADF

NO. 1 ADF/VOR PUSHBUTTON SELECTOR— SELECTS ALTERNATIVE NO. 1 ADF OR VOR MODE MODE INDICATOR—

NO. 2 ADF/VOR PUSHBUTTON SELECTOR— SELECTS ALTERNATIVE NO. 2 ADF OR VOR MODE

DIRECTION OF ARROW INDICATES MODE SELECTED

Figure 22-19. Radio Magnetic Indicator

22-28

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

RADIO MAGNETIC INDICATOR (RMI)

NOTES

The RMI is a dual-pointer, rotating, servodriven azimuth dial instrument which shows airplane heading under the lubber line. A single- bar pointer and a wide, double-bar pointer provide bearing indication for either ADF or VOR modes by means of rotary switches or “push’’ transfer buttons (Figure 22-19), depending on the particular instrument. Servo error, compass valid, and instrument power are monitored by a single power-off or heading flag, depending on the particular instrument. The head of the bearing pointer read against the rotating compass card shows the magnetic bearing of the ground station from magnetic north; the number of degrees clockwise from the magnetic heading of the airplane to the bearing pointer shows the relative bearing of the station from the airplane. Magnetic heading (compass) information is from the cross-side AHRS.

Revision 4

22-29

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SEL

KIAS H

20

SEL

KIAS

ch 200 -6

20

H

20

20

60

10

10

20

10

10

ch 200 -6

-2

60

10

10

-2

-1 0

10

20

-1

EADI

EADI

-1

0

-1

10

-2 80

CRS

20

HDG FAIL

CMP1

33 30

-2

-6 FPM

20

0

WEATHER RECEIVER/ TRANSMITTER AND ANTENNA

NAV1 2.1 NM

CRS

-6 FPM

20

0

HDG FAIL

CMP1

33 30

NAV1 2.1 NM

N

24 S

15

12 VOR 1 OFF OFF ADI DIM

ATT REV

HDG REV RNV 2 ADF 2

OFF DH TST

HSI DIM

WX DIM

DISPLAY CONTROLLER

NAV REV

FULL ARC

WEATHER DISPLAY

AUX 2

VOR 2 OFF

BRG

AUX 1

BRG

WX

RNV 1 ADF 1 VOR 1 OFF OFF ADI DIM

GSPD TTG

MAP

ATT REV

HDG REV

NAV REV

RNV 2 ADF 2

TEST

15

S

BRG

RNV 1 ADF 1

GSPD TTG

MAP

TEST

AUX 1

WX

GSPD 130 KTS

HDG

DISPLAY CONTROLLER FULL ARC

12

GSPD 130 KTS

HDG

V

E

21

6

E

21

EHSI

EHSI

V

6

24

3

3

W

W

N

20

80

OFF DH TST

HSI DIM

WX DIM

AUX 2

VOR 2 OFF

BRG

RIGHT SENSORS LEFT SENSORS COURSE

SYMBOL GENERATOR ASCB

HEADING

INSTRUMENT REMOTE CONTROLLERS

COURSE HEADING

ASCB

SYMBOL GENERATOR

Figure 22-20. Sperry EFIS Data Flow

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Revision 4

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) GENERAL The Sperry EDZ-811 electronic flight instrument system has the following components: • Four electronic display units (two EADIs and two EHSIs)

The symbol generators interface with the ASCB to provide the required outputs to the flight instruments and for flight control purposes. Input data is from the following sources: • VOR/localizer/glide-slope receiver or MLS • Attitude and heading reference system (AHRS) • Radio altimeter

• Two symbol generators

• Air data system

• Two display controllers

• Distance-measuring system

• Two instrument remote controllers

• Flight control system

The EFIS displays attitude, compass, and navigational information on the electronic display units using color cathode ray tubes (CRTs). In addition to information previously displayed on electromechanical ADIs and HSIs, facilities such as air data, waypoint and VOR/localizer locations, and weather patterns are also available.

• Long-range navigation system (LRN, FMS, INS, etc.)

The Sperry EFIS uses four identical electronic flight display units together with the necessary processor units and control panels. Figure 22-20 shows a simplified block diagram of the system.

• Angle-of-attack system

Two identical electronic flight display units are installed one above the other on each main instrument panel. These correspond to the conventional placement of electromechanical ADIs and HSIs and are referred to as the EADIs and EHSIs. A standard ball-in-fluid inclin ometer must be mounted on the CRT that is to be used as an EADI. Two (pilot’s and copilot’s) display controllers provide the means by which the pilots can control the display formatting, such as full compass or partial compass (sector scan) display.

Revision 2

• Vertical navigation system • Weather radar system • Automatic direction-finding system

Output data include display steering commands and other navigational data for flight control purposes as well as information of an advisory nature. Extensive monitoring and comparator circuits provide warning flags and other types of data flagging to indicate possible equipment malfunctions. This chapter presents a general description and operation of EFIS units. For details of each display and its controls, consult the Flight Manual Supplement, Maintenance Manual, and the vendor manual for a specific installation. Figure 22-21 shows several typical EFIS displays. Sheet 2 of 2 is on the following page.

22-31

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NORMAL ATTITUDE DISPLAY

COMPOSITE DISPLAY Figure 22-21. Sperry EFIS Displays (Typical) (Sheet 1 of 2)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PLAN MODE DISPLAY

PARTIAL COMPASS DISPLAY

Figure 22-21. Sperry EFIS Displays (Typical) (Sheet 2 of 2)

Revision 2

22-33

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ATTITUDE SOURCE ROLL ANNUNCIATOR ATTITUDE (ATT 1 OR ATT 2) POINTER

ROLL SCALE

ATTITUDE SPHERE GO AROUND ANNUNCIATOR GA ATT2 AOA F

PITCH AND ROLL COMMAND POINTERS

20

20 10

10

GLIDE-SLOPE (G) VNAV (V), OR ELEVATION (E) DEVIATION POINTER

G

FAST/SLOW POINTER

10

S 20 200 DH

10 20

AIRCRAFT SYMBOL

I

DH

MARKER BEACON 0 (BLUE) M (AMBER) I (WHITE)

140 RA

DECISION HEIGHT DISPLAY EXPANDED INCLINOMETER RISING LOCALIZER OR RUNWAY AZIMUTH POINTE (NOTE)

RADIO DECISION HEIGHT ALTITUDE DISPLAY (AMBER)

NOTE: WHEN NOT TUNED TO AN ILS FREQUENCY, THE RATE OF TURN POINTER AND SCALE IS PRESENT IN PLACE OF THE EXPANDED LOCALIZER. WHEN MLS IS SELECTED, THE EXPANDED LOCALIZER POINTER DISPLAYS AZIMUTH DEVIATION.

Figure 22-22. EADI Symbology COMPARATOR MONITOR CAUTION (AMBER) FAST/SLOW FAILURE (RED)

ATT FAIL ANNUNCIATOR (RED)

GLIDE-SLOPE VNAV, OR ELEVATION FAILURE (RED)

SG2 AOA F

HDG

FLIGHT DIRECTOR FAILURE (AMBER)

SAME ATTITUDE SOURCE (AMBER)

FD FAIL ATT2 20 AOA

PIT ROL ATT

LOC GS ILS AZ EL MLS

HDG ATT ILS

20

F

ATT FAIL

10

10 G

10

10

S

S

20

M DH

DH 140 RA

200 DH

EXPANDED LOCALIZER OR AZIMUTH FAILURE (RED) *

RED

20

COMMON SYMBOL GENERATOR ANNUNCIATION (AMBER)

M DH

DH FAILURE (AMBER DASHES)

RA

DECISION HEIGHT WARNING (AMBER)

RA FAILURE (AMBER DASHES)

AMBER

Figure 22-23. EADI Caution and Failure Annunciations

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EADI

NOTES

The EADI (Figure 22-22) combines the familiar attitude sphere display with lateral and vertical computed steering signals to provide the pilot commands required to intercept and maintain a desired flight path. Figure 22-23 shows EADI caution and failure annunciations. The EADI provides the following display information: • Glide-slope or elevation deviation • Expanded localizer or azimuth deviation • Radio altitude • Rising runway • Digital • Decision height • Marker beacon annunciation • Rate of turn • Speed command • Attitude source Other items which can be displayed on the EADI, depending upon pilot selection or operational mode of the airplane, include the following: • Airspeed and airspeed trend • Attitude comparator warnings • Decision height set

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DRIFT HEADING ANGLE SOURCE BUG ANNUNCIATOR

HEADING SELECT BUG

FORE LUBBER LINE

BEARING POINTERS

HEADING DIAL

COURSE / DESIRED TRACK DISPLAY CRS

315

H

COURSE SELECT POINTER

NAVIGATION SOURCE ANNUNCIATOR

NAV1

CMP1

33

30

2.1 NM

N

DISTANCE DISPLAY

W

WX

DME HOLD

3

24

WEATHER RADAR MODE ANNUNCIATOR

VERTICAL NAVIGATION GLIDE-SLOPE, OR ELEVATION DEVIATION POINTER

E

21

6

15

S ADF HDG

12

VOR1

TO-FROM ANNUNCIATOR

G

GSPD 130 KTS

319

V.G. OR E. ANNUNCIATOR

BEARING POINTER SOURCE ANNUNCIATOR

GROUND SPEED DISPLAY (NOTE)

HEADING SELECT DISPLAY

AIRCRAFT SYMBOL

AFT LUBBER LINE

RECIPROCAL COURSE POINTER

NAV SOURCE ANNUNCIATOR

NOTE: TIME-TO-GO IS ALSO DISPLAYED AT THIS LOCATION. HEADING SOURCE ANNUNCIATOR

CRS OR DTRK

COURSE OR AZIMUTH DEVIATION BAR

HDG1 HDG2 MAG1 MAG2

NAV1 NAV2 MLS1 MLS2 VLF VLF1 VLF2 VLF3 RNV

TARGET ALERT

DG1 DG2

VAR (AMBER) TGT (AMBER) TGT (GREEN)

RNV1 RNV2 INS INS1 INS2 INS3 FMS1 FMS2

WEATHER RADAR MODE ANNUNCIATOR

WX

33

VERTICAL DEVIATION ANNUNCIATOR V G E

N

W

30

24 21 S

ADF HDG

319

15

VOR1

G

12

INS INS1 INS2 FMS1 FMS2

315

E

ADF1 ADF2 VOR1 VOR2 RNV1 RNV2

TGT NAV1 2.1 NM

CMP1

6

BEARING SOURCES

CRS

3

WAIT (GREEN) STBY (GREEN) WX (GREEN OR AMBER) GMAP (GREEN) TEST (GREEN) FAIL (AMBER) RCT (GREEN) GCR (AMBER) CR/R (AMBER) CYC (GREEN)

GSPD 130 KTS

TIME-TO-GO OR GROUND SPEED DISPLAY TTG 399 MIN GSPD 999 KTS

FULL COMPASS FORMATS

Figure 22-24. EHSI Symbology (Sheet 1 of 2)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

CMP1

DTRK

32

FMS1

315

30 NM

NAVAID

33

30

RANGE RINGS

N W

G RANGE ANNUNCIATION (NAUTICAL MILES) * 5 12.5 25

VOR1 ADF HDG

50 100 150

GSPD 260 KTS

25 15

319

* RANGE ANNUNCIATION ON INNER RING IS 1/2 THE RANGE SETTING OF THE WEATHER RADAR. WIND VECTOR DISPLAY

TO-FROM HEADING SOURCE ANNUNCIATOR ANNUNCIATOR

HEADING DISPLAY

NAVIGATION SOURCE ANNUNCIATOR

COURSE/DESIRED TRACK DISPLAY

DISTANCE DISPLAY DTRK

HEADING SELECT BUG

315

CMP1 TO 30

HEADING DIAL

32

FMS1 30 NM

05 33

RANGE RINGS

N W

VOR STATION NAVAID, BLUE FOR VOR 1 GREEN FOR VOR 2

04

WEATHER 03 50

MULTIPLE LRN WAYPOINTS

VOR2 HDG

GSPD 260 KTS

319

AIRCRAFT SYMBOL GROUND SPEED DISPLAY (NOTE)

NAVAID SOURCE ANNUNCIATOR

HEADING SELECT DISPLAY

COURSE DEVIATION BAR AND SCALE

NOTE: TIME-TO-GO IS ALSO DISPLAYED AT THIS LOCATION.

PARTIAL COMPASS FORMATS

Figure 22-24. EHSI Symbology (Sheet 2 of 2)

Revision 4

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TO-FROM ANNUNCIATOR

ATTITUDE ROLL SOURCE ATTITUDE ANNUNCIATOR POINTER

NAVIGATION ROLL DISTANCE SOURCE SCALE ANNUNCIATOR DISPLAY

COURSE/DESIRED TRACK DISPLAY

ATTITUDE SPHERE

AIRCRAFT SYMBOL

CRS FR

HEADING DISPLAY

022

NAV1 ATT2

20

HEADING SELECT DISPLAY

PITCH AND ROLL COMMAND BARS

120 NM

20 10

VERTICAL NAVIGATION, GLIDE SLOPE, OR ELEVATION DEVIATION POINTER AND SCALE

10 G

10

HEADING SOURCE ANNUNCIATOR

010 CMP 1

10

COURSE SELECT POINTER MARKER BEACON

000 3 3

0 0

DH

200 DH

I

0 3

140 RA

RADIO ALTITUDE DISPLAY

DECISION HEIGHT DISPLAY COURSE OR COURSE OR DECISION FORE AZIMUTH AZIMUTH HEIGHT LUBBER DEVIATION DEVIATION BAR LINE DISPLAY

HEADING TAPE DISPLAY

Figure 22-25. EADI/EHSI Composite Display HEADING FAILURE DISPLAY (RED)

VERTICAL DEVIATION ANNUNCIATOR (WHITE)

HEADING SELECT FAILURE (AMBER DASHES)

RED

NOTES: 1. GLIDESLOPE OR ELEVATION DEVIATION FAILURE IS SIMILAR. 2. IN THE EVENT OF HEADING FAILURE, THE COURSE SCALE AND RED X WILL NOT BE DISPLAYED AND THE CRS AND HDG READOUTS WILL INDICATE AMBER DASHES.

WPT VOR1 ADF HDG

33 N

24 W

30

WAYPOINT ALERT ANNUNCIATOR (AMBER)

V

GSPD 130 KTS

CRS CMP1

21

21 12 15 S

HDG

VERTICAL DEVIATION FAILURE (RED) (NOTE 1)

TGTNAV1 NM H

G

6 E

6 E

AMBER DASHES (NOTE 2)

NAV1 2.1 NM

WEATHER TARGET ALERT (AMBER)

3

24 W

HDG CRS CMP1 FAIL 33 N 30 3

COURSE OR AZIMUTH DEVIATION FAILURE (RED) (NOTE 2)

COURSE SELECT FAILURE (AMBER DASHES) (NOTE 2)

SAME HEADING OR NAV SOURCE (AMBER) (NOTE 3)

12 15 S

AMBER DASHES (NOTE 2)

DME DISPLAY FAILURE (AMBER DASHES) (NOTE 4) DME HOLD ANNUNCIATOR (AMBER)

GSPD KTS

AMBER

GROUND SPEED FAILURE (AMBER DASHES) (NOTE 1)

NOTES: 1. TTG FAILURE IS SIMILAR. 2. DTRK FAILURE IS SIMILAR. 3. IF THE PILOT SELECTS NAV 2 AND THE COPILOT SELECTS NAV 1, THE NAV ANNUNCIATORS ARE AMBER. 4. AN AMBER N/A IS DISPLAYED WHEN SELECTED LRN DISTANCE IS NOT AVAILABLE.

Figure 22-26. EHSI Caution and Failure Annunciations

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EHSI The EHSI combines numerous displays to provide a map-like presentation of the airplane pos i t i o n . T h e i n d i c a t o r s h ow s a i r p l a n e displacement relative to VOR radials, localizers, and glide-slope beams. At power-up, the EHSI presents a full compass display. By pressing the FULL/ARC alternate-action pushbutton switch on the display controller, the full compass display is changed to a partial compass format (sector scan). If weather radar returns are desired, pressing the WX button on the display controller also changes the full compass display to a partial compass format displaying weather radar returns or adds weather returns to the display if already in partial compass display format. This also happens in MAP function when the alternate action MAP button is depressed.

The following information is available in partial compass displays: • Weather radar • Wind vector • Navigation map (range annunciation and waypoints) (KNS 660, optional) • Multiple waypoints (KNS 660, optional) Figure 22-24 illustrates the EHSI symbology in the full and partial compass formats. Figures 2 2 - 2 5 a n d 2 2 - 2 6 s h ow a n E A D I / E H S I composite display and describe the EHSI caution and failure annunciations, respectively.

NOTES

The following information is available in full compass displays: • Heading • Course select • Course or azimuth deviation • Distance • Groundspeed • To/from • Desired track • Bearings 1 and 2 • Heading select • Vertical, glide-slope, or elevation deviation • Time to go • Weather mode annunciator • Heading sync

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-27. EFIS Controller DME HOLD

INN OUT MID

TAS

REVN

SG FAIL

WHEN A SYMBOL GENERATOR FAILURE OCCURS, THIS IS THE SWITCH USED TO RESTORE THE DISPLAY USING THE CROSSSIDE SYMBOL GENERATOR.

*

* SG FAIL IS AMBER ON THE EHSI. INTERNAL SYSTEM FAILURE

Figure 22-28. Symbol Generator Failure Indication

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EFIS CONTROLLER

• DH—Allows setting of decision height

Two EFIS controllers (Figure 22-27) are located on the center console and allow selection of the desired display format on the EFIS display units. Following are the controls and their functions:

• TEST—Allows system test

• Bearing source select knob markings:

BEARING 1 BRG OFF VOR1 ADF1 RNV1 AUX1

°

BEARING 2 BRG ◊ OFF VOR2 ADF2 RNV2 AUX2

• FULL/ARC—Alternates between full and partial compass displays • WX—Allows weather to be displayed on a partial compass display or, if in full compass format, changes the display to partial format with weather information

INSTRUMENT REMOTE CONTROLLER Two instrument remote controllers are located on the glareshield, one for the pilot's instruments and one for the copilot's instruments. On electromechanical flight instrument installations, the controllers are connected to the HSIs. On EFIS-equipped airplanes, the controllers are connected to the symbol generators. With either installation, the HEADING knob sets the HSI heading marker (bug), and the COURSE knob sets the HSI course arrow. Digital readouts are also shown in the upper-left corners of the HSIs for course, and in the lower-left corners for heading.

SYSTEM TEST

• MAP—Changes a full compass display to a partial compass format, allowing one waypoint for each bearing pointer and VOR/DME ground station position to be displayed if the station is within the range selected

Pressing the TEST button causes flags and cautions to be displayed and checks the radio altimeter. EFIS test is functional on the ground only. The radio altimeter check is functional at all times except during glide-slope capture/track and expanded localizer capture/track. When the TEST button is pressed:

• GSPD/TTG—Alternates between displaying groundspeed or time to go in the lower-right corner of the EHSI

• The radio altimeter reads the test value (slews to 100 feet for Sperry radio altimeter).

• AT T R E V, H D G R E V, NAV R E V — Allows cross-side attitude information, heading, or navigation source information to be displayed on the EHSI

• All flags are in view as indicated by a red X through all pointer scales.

• DIM—Two inputs contribute to the overall display brightness:

• The word TEST appears in the upper-left center of the EADI.

• Ambient light sensed by the photosensors on each EFIS display • Manual setting of the DIM controls

Revision 2

• Command cue is out of view.

NOTE Self-test is inhibited during glideslope capture.

22-41

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DIM SWITCH THREE-POSITION SWITCH CONTROLS DISPLAY BRIGHTNESS AND TIME ZONE CORRECTION: B—BRIGHT DISPLAY FOR DAYLIGHT OPERATION DIM—DIM DISPLAY FOR NIGHT OPERATION 1 HR UP (MOMENTARY)—ADVANCES CLOCK TIME BY ONE HOUR TO ADJUST FOR TIME ZONES

D DIM

SET

1 HR UP

1 2 5 9 TIME

HOURS DISPLAY SET SWITCH ADJUSTS CLOCK TIME SETTING: UP (MOMENTARY)—RUNS CLOCK AT DOUBLE SPEED TO ADVANCE TIME SETTING BY ONE SECOND FOR EVERY SECOND HELD D (MOMENTARY)—STOPS CLOCK TO SET BACK TIME SETTING BY ONE SECOND FOR EVERY SECOND HELD DISPLAY SELECT SWITCH THREE-POSITION SWITCH SELECTS MODE TO BE DISPLAYED: TIME—CLOCK TIME F.T.—FLIGHT TIME E.T.—ELAPSED TIME

F. T. E.T.

DAVTRON

5 9

ZERO S T O P R

M811B

SECONDS DISPLAY ELAPSED TIME SWITCH THREE-POSITION SWITCH CONTROLS ELAPSED TIME FUNCTION: RUN—STARTS ELAPSED TIME STOP—STOPS ELAPSED TIME ZERO (MOMENTARY)—CLEARS ELAPSED TIME

Figure 22-29. Davtron Clock

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

MISCELLANEOUS AND STANDBY FLIGHT INSTRUMENTS

NOTES

CLOCK Two Davtron digital electronic clocks (Figure 22-29) are located in the glareshield. Each provides three independent functions: clock time (TIME), flight time (FT), and elapsed time (ET). These functions are displayed on the hours, minutes, and seconds display by selecting the display select switch to the desired function. All functions continue to operate normally regardless of which function is selected for display. Internal functions of the clock are powered by an internal battery: however, the display is powered by airplane 28-VDC power. To erase flight time with the aircraft on the ground and power applied, a flight time erase button is located on the bottom of the clock, accessible underneath the glareshield.

Revision 2

22-43

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

LUBBER LINE

15 12

COMPASS CARD

N-S

E

6

E-W

Figure 22-30. Standby Magnetic Compass

9 8 7

0 1 0 000

1 2

ALT

3

IN HG

6

2992 5

4

Figure 22-31. Standby Altimeter

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

STANDBY MAGNETIC COMPASS

NOTES

A standby magnetic compass (Figure 22-30) is located at the top of the windshield center post. Compass heading is read by the lubber line against the compass card. A post light is provided. Care must be taken if the standby compass is used because the compass deviation (magnetic heading ± deviation = compass heading) can vary and may be unreliable due to the amount of electrical equipment which is energized in the cockpit. Windshield heat may cause erratic and unreliable compass readings.

STANDBY ALTIMETER A standby altimeter (Figure 22-31) is located on the center instrument panel. Altitude is displayed in the range of –1,000 to 50,000 feet by a digital counter and a pointer read against a dial. The pointer makes one revolution of the dial per 1,000 feet. Gradations in 20-foot increments are numbered at 100-foot intervals. The counter shows ten-thousands and thousands of feet. A knob is provided to set the baroscale counter which simultaneously corrects the setting of the pointer and digital counter. The baroscale counter is usually read in inches of mercury, but instruments reading in millibars are available. The instrument is connected to the pilot’s static air ports.

Revision 2

22-45

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-32. Standby Attitude Indicator

STANDARD RATE TURN INDEXES

TURN NEEDLE

OFF FLAG

INCLINOMETER

Figure 22-33. Turn-and-Slip Indicator

22-46

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

STANDBY ATTITUDE INDICATOR

NOTES

The standby attitude indicator (Figure 22-32) is located on the pilot’s instrument panel. It is gyro-stabilized and requires 28-VDC power. A striped “off’’ flag indicates an invalid display due to inadequate electrical power or gyro speed. A caging knob is located in the lower-right corner of the instrument.

TURN-AND-SLIP INDICATOR The turn-and-slip indicator (Figure 22-33) provides basic “needle and ball’’ turn rate and slip information. The indicator needle is stabilized by an integral electric gyro. Deflection to the index mark on either side of center represents a standard-rate turn of 180° per minute. A red flag appears in each indicator when it is deenergized (supply power falls below 23.4 VDC). The turn-and-slip indicator is optional with E/M instruments, and is not installed on EFIS-equipped aircraft.

Revision 2

22-47

22-48 AD-550C ADI GA

AD-550C ADI GA

DH FD

10

10

10

10

20

20

1570

DH

RAD ALT

DIM

9 0 1 8 28 720 2 ALT IN HG 3 MB

7

DH SET/ TEST

1013

6 BARO

WI-800 INDICATOR

2992

5

4

9 0 1 8 28 720 2 ALT IN HG 3 MB

7

1013

6 BARO

10 20

1570

DH

5

RAD ALT

DIM

DH SET/ TEST

4

RI-306-IRC

33

W

N A V

3

E

E

Al-801 ALTITUDE PRESELECT CONTROLLER

V E R T

6

6

V E R T

24

12

12

24

DIST

30

3

00 0 COURSE

C

WR-800 RECEIVER/TRANSMITTER

HD

DIST

30

33

N A V

10

200

RD-550A HSI WA-800 ANTENNA

C

00 0 COURSE

10

20

SLOW

ATT 20

10

2992

RD-5504 HSI HD

RI-306-IRC

DH

20 FAST

HLD NAV PRE 12 MLS

15

21

SEL

OFF

SERIAL DATA

SLOW

SLOW

FAST

DG

FAST

HDG

VG ERECT

DG

21

NM

KTS MIN

PWR

AUX PR OFF TEST

SLOW

FAST

SLOW

FAST

DG SLEW

DG

FAST

HDG

VG ERECT

DG

AUX PR

TEST

FAST

DG SLEW SLAVE

BASIC

SLAVE

BASIC

SERIAL DATA

AC-801 AHRS CONTROLLER/ REMOTE COMPENSATOR'S

NAV SWITCHING

FX- 600 THIN FLUX VALVE

AH-600 STRAPDOWN AHRU

AH-600 STRAPDOWN AHRU

FX- 600 THIN FLUX VALVE

DUAL ASCB DATA BUSSES WHITE AMBER WHITE GREEN

AZ-810 AIR DATA COMPUTER

SERIAL DATA FZ-800 FLIGHT GUIDANCE COMPUTER

NAV SWITCHING AZ-810 AIR DATA COMPUTER

WHITE AMBER WHITE GREEN

ID-802 ADVISORY DISPLAY

ID-802 ADVISORY DISPLAY

VOR/LOC,MLS, RNAV,AUX NAV, RCVRS

HDG

BANK

ALT

AP

NAV

CAT 2

VS

YD

APP

STBY

FLC

M TRIM

B/C

TINY

VNAV

CPL

VOR/LOC,MLS, RNAV,AUX NAV, RCVRS

SERIAL DATA GC-801 FLIGHT GUIDANCE CONTROLLER

FZ-800 FLIGHT GUIDANCE COMPUTER

RT-300 RADIO ALTIMETER RECEIVER/ TRANSMITTER

SM-300 ELEVATOR SERVO AT-300 ANTENNA

15

888 8 8

AT-300 ANTENNA

RT-300 RADIO ALTIMETER RECEIVER/ TRANSMITTER

SM-300 AILERON SERVO

SM-710 LINEAR ACTUATOR RUDDER YAW DAMPER

TM-400 ELEVATOR TRIM SERVO AT-300 ANTENNA

AT-300 ANTENNA

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

200

BA-141 ALT IND

W

SLOW

BA-141 ALT IND

20

20 FAST

FD

ATT

Canada Ltd.

Revision 4

Figure 22-34. Sperry DFZ-8000—Electromechanical Flight Instruments

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

SPERRY DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM

NOTES

GENERAL The Sperry SPZ-8000 digital integrated flight control system is a complete automatic flight control system which provides fail-operational execution of flight director guidance, autopilot, yaw damper, and trim functions. Major components include the air data system, the auto pilot and flight director, and either the standard electromechanical ADIs and HSIs or the EFIS. Specifically, the SPZ-8000 subsystems consist of the: • Attitude and heading reference system (AHRS) • Air data system • Radio altimeter system • Standard flight instrument system • Electronic flight instrument system (optional) • Dual flight guidance system • Advisory display This chapter gives a general description and operation of the above systems. The user should consult the AFM, Flight Manual Supplement, Maintenance Manual, and vendor manual for complete description and operation of the systems installed on specific airplanes and optional equipment. To assist in clarifying how the system operates, two simplified functional block diagrams show how the system components are interconnected. Figure 22-34 s h ow s t h e s y s t e m w i t h s t a n d a r d f l i g h t instruments installed.

Revision 2

22-49

22-50 ED-800 EADI

BA-141 ALT IND DC-810 DISPLAY CONTROLLER

CMP 1

3 3

MAP

RNV 1 ADF 1

GSPD TTG

ATT REV

HDG REV

NAV REV

8 28 7

10

000

010

WX

AUX 1

0 0

I

0 3

VOR 1 OFF OFF ADI DIM

RNV 2 ADF 2 OFF DH TST

HSI DIM

WX DIM

1013

6

AUX 2 BARO

VOR 2 OFF

8 28

WI-800 INDICATOR

2992

5

4

7

MB

1013

6 BARO

720

2 ALT IN HG 3 2992

5

BRG

ED-800 EHSI

WA-800 ANTENNA HDG FAIL

CMP1

30

ATT REV

VOR 1 OFF OFF ADI DIM

HSI DIM

W

21

12

SEL

SERIAL

15

S

RI-306-IRC

OFF SLOW

SERIAL DATA

SLOW

FAST

DG

FAST

HDG

VG ERECT

DG

PWR

AUX PR OFF TEST

SLOW

FAST

SLOW

FAST

DG

FAST

HDG

VG ERECT

DG

CMP 1

3 3

0 0

0 3 DH

SLAVE

FX 600 THIN FLUX VALVE

AH-600 STRAPDOWN AHRU

SERIAL DATA FZ-800 FLIGHT GUIEANCE COMPUTER

AT-300 ANTENNA

NAV1 2.1 NM

N

V

GSPD 130 KTS

AH-600 STRAPDOWN AHRU NAV SWITCHING AZ-810 AIR DATA COMPUTER

WHITE AMBER WHITE GREEN

ID-802 ADVISORY DISPLAY HDG

BANK

NAV

CAT 2

VS

YD

APP

STBY

FLC

ALT

M TRIM

B/C

TINY

VNAV

CPL

VOR/LOC,MLS, RNAV,AUX NAV, RCVRS

AP

SERIAL DATA GC-801 FLIGHT GUIDANCE CONTROLLER

FZ-800 FLIGHT GUIDANCE COMPUTER RT-300 RADIO ALTIMETER RECEIVER/ TRANSMITTER

RT-300 RADIO ALTIMETER RECEIVER/ TRANSMITTER SM-300 ELEVATOR SERVO

33

RI-306-IRC

SERIAL DATA

DUAL ASCB DATA BUSSES

ID-802 ADVISORY DISPLAY

30

HDG

AUX PR

NAV SWITCHING AZ-810 AIR DATA COMPUTER

HDG FAIL

CMP1

BASIC

FX 600 THIN FLUX VALVE

WHITE AMBER WHITE GREEN

I

140 RA

ED-800 EHSI CRS

AC-801 AHRS CONTROLLER/ REMOTE COMPENSATOR

AT-300 ANTENNA

10

000

010

200 DH

BRG

TEST

FAST

DG SLEW BASIC

TO PILOT'S EADI

G

AUX 2

VOR 2 OFF

SG-311 SYMBOL GENERATOR

KTS MIN

DG SLEW SLAVE

VOR/LOC,MLS, RNAV,AUX NAV, RCVRS

10

E

E

888 8 8

SG-311 SYMBOL GENERATOR

10

6

6

V

NM

GSPD 130 KTS

WX DIM

120 NM

20

NAV REV

RNV 2 ADF 2 OFF DH TST

HLD NAV PRE 12 MLS

HDG

HDG REV

20

10 AUX 1

AL-801 ALTITUDE PRESELECT CONTROLLER

3

24

RNV 1 ADF 1

GSPD TTG

WR 800 RECEIVER/TRANSMITTER

NAV1 2.1 NM

N

MAP

NAV1 ATT2

3

TO COPILOT'S EHSI

33

WX

022

SERIAL

CRS

4

FULL ARC

BRG

SERIAL

BRG

720

2 ALT IN HG 3

CRS FR

SM-300 AILERON SERVO

SM-710 LINEAR ACTUATOR RUDDER YAW DAMPER

TM-400 ELEVATOR TRIM SERVO AT-300 ANTENNA

AT-300 ANTENNA

TO PILOT'S EHSI

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

140 RA

DH

200 DH

MB

DC-810 DISPLAY CONTROLLER

SERIAL

10

FULL ARC

1

TEST

G

0

9

W

10

ED-800 EADI

BA-141 ALT IND

1

24

10

0

9

21

120 NM

20

S

20

TEST

TO COPILOT'S EADI

NAV1 ATT2

15

022

12

CRS FR

Canada Ltd.

Revision 4

Figure 22-35. Sperry DFZ-8000—Electronic Flight Instruments

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS)

Autopilot Modes and Controls

General

Following are the autopilot basic operating modes:

The Sperry DFZ-8000 automatic flight control system provides flight director, autopilot, and yaw damper functions. The system design is fail-operational for the first failure of any system component with the exception of the autopilot servoactuators; a single actuator is used in each axis due to the high reliability of these units. Two flight guidance computers provide fail-operational capability. In normal operation, one flight guidance computer is active, operating in the normal fail-operational mode; the second computer is doing all the same processing as the active computer but acts as a hot spare. If the active computer fails and disengages, the second computer automatically picks up where the first computer failed. An advisory is then displayed on the advisory display as to why the automatic switchover was required. Figures 22-34 and 22-35 illustrate the digital flight guidance system. As shown in these figures, the AFCS receives information from the dual attitude and heading reference systems, the dual ADCs, the ADI and HSI, the navigation radios, and the flight management system (if installed). The interconnection of these systems is provided by the bidirectional avionics standard communications bus (ASCB). Because the bus is controlled by the flight guidance computer, failure of a flight guidance computer or the bus controller function within a flight guidance computer usually but not always, depending on the malfunction, causes the bus control to be automatically transferred to the remaining computer.

Autopilot Modes

• Heading hold—When the autopilot is engaged and no lateral mode is selected on the AFCS controller, the autopilot is in basic heading-hold mode. If the autopilot is engaged while the airplane is in a bank, the system engages wings level; once the bank angle is less than 3° for ten seconds, the heading-hold mode is engaged (not annunciated on the advisory display). • Roll hold—The autopilot holds the desired roll attitude when the touch control steering (TCS) button is released between 6 and 45° of roll attitude. If the button is released when roll attitude is less than 6°, the airplane rolls to and flies wings-level. If the button is released when roll attitude is greater than 45°, the airplane rolls back to 45° and holds that banked attitude. The TCS mode causes “TCS ENGAGE’’ to be displayed on the ID 802 advisory display, and the AP engage arrow on the AFCS controller to be extinguished, indicating that the autopilot servos are not engaged.

NOTE To enter roll hold, the following conditions must be met: • Autopilot engaged • No lateral mode engaged • TCS used to initiate the roll (Roll hold is not annunciated on the advisory display.)

Revision 2

22-51

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-36. AFCS Controller

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

• Pitch hold—When the autopilot is engaged and no vertical mode is selected on the AFCS controller, the autopilot is in basic pitch-hold mode. If a lateral flight director mode has been selected, but no vertical mode, the flight director engages the pitch-hold mode, and a pitch command bar is displayed. Pitch angle limit using the pitch wheel is ± 20° (pitch is annunciated on the advisory display).

Autopilot Controls General The flight guidance controller (AFCS controller) is used to engage and disengage the auto pilot and yaw damper, as well as selecting flight director modes of operation. Most of the controls on the AFCS controller (Figure 22-36) are alternate-action pushbuttons (push on, push off).

To engage the autopilot on the ground for maint e n a n c e p u r p o s e s , t h e f o l l ow i n g c i r c u i t breakers must be pulled to bypass the weighton-wheels circuits: • Right essential circuit-breaker panel— CB M6 • Right main circuit-breaker panel— CB N6 • Left essential circuit-breaker panel— CB E6 • Left main circuit-breaker panel—CB F6

NOTE Use this procedure to perform ground maintenance testing only.

NOTES

The AP and YD buttons have left and right engage arrows which indicate system engagement when illuminated and from which flight director (Number 1, left or Number 2, right) the autopilot is receiving flight guidance information. The Number 1 (left) FGC is automatically selected during initial power-up. If desired, selection of the Number 2 (right) FGC (AP and YD buttons) is made with the R AFCS button on the advisory display. Selection of L AFCS returns control to the left side.

Autopilot Engagement and Disengagement Activation of the AP pushbutton engages the autopilot and yaw damper functions simultaneously. A repeated activation of the AP button disengages only the autopilot function.

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Figure 22-37. GA Buttons

Figure 22-38. Control Wheel Switches

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Autopilot disengagement occurs from any of the following actions: • Pressing the AP or YD pushbuttons on the AFCS controller • Pressing either GA button (Figure 22-37) on the power levers • Pressing the AP DIS (disconnect) switch on either control wheel (Figure 22-38) • Pressing the stick shaker self-test button • Activation of the stick shaker

Yaw Damper Engagement and Disengagement

IAS hold or VS hold mode is engaged, the wheel is used to dial in a new reference, which is displayed on the advisory display as part of the mode message.

Touch Control Steering When depressed, the touch control steering (TCS) button on the control wheels uncouples the autopilot servos from the airplane without disengaging the auto pilot. When the TCS button is released, the autopilot holds the current pitch attitude of the airplane. When an air data mode (i.e., IAS, VS, or ALT hold) is engaged, the TCS button can be used to synchronize the AFCS command to a selected reference.

Flight Director Modes and Controls

Activation of the YD pushbutton engages the yaw damper; a second push disengages it. The yaw damper system is independent of the autopilot, but an electrical interlock ensures that the autopilot cannot be engaged without also engaging the yaw damper. Disengaging the yaw damper also disengages the autopilot, if engaged.

Flight guidance consists of lateral and vertical modes as selected on the AFCS controller. The following mode capabilities are available.

HSI SEL

Lateral Modes

The HSI SEL (select) pushbutton on the AFCS controller is used to select either the pilot’s or copilot’s HSI information. The appropriate arrow annunciator on either side of the pushbutton illumi nates, indicating from which HSI and ADC the autopilot is receiving information.

Heading Select (HDG) When HDG is selected on the AFCS controller, the airplane will roll to follow the heading bug on the active HSI. HDG is annunciated on the advisory display. HDG may be canceled by:

NAV SEL The NAV SEL pushbutton alternately allows either the pilot or copilot to select the desired navigation source (V/L, MLS, or AUX) to be fed to the HSI for navigation.

Pitch Wheel When the autopilot is in basic pitch-hold mode, moving the thumbwheel toward NOSE UP or NOSE DN causes an appropriate change in pitch attitude at a rate proportional to the amount of pitch wheel displacement. When

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General

• Pressing HDG on the AFCS controller • Pressing STBY on the AFCS controller • Selecting go-around • Automatic capture of any other lateral steering mode • Coupling of the cross-side HSI

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Navigation (NAV) When NAV is selected, flight guidance roll commands to the ADIs or EADIs and the flight control function are computed to command the airplane to intercept, capture, and track whatever navigation problem is being displayed on the active HSI. The VOR/ILS mode has an armand-capture phase which allows a path intercept and automatic transition into navigation capture and track. NAV provides localizer approach if glide-slope tracking is not desired. Composite steering signals are used for FMS tracking. The raw data is displayed on the HSI. When this mode is selected, the flight guidance computer arms until directed by the FMS to capture and track steering commands. Approach (APP) Lateral approach mode is selected by depressing the APP button on the AFCS controller. Flight guidance commands and attitude commands to the flight control function are computed to intercept, capture, and track whatever navigation path is being displayed on the active HSI. This allows optimized ILS, VOR, and RNAV approach mode lateral control, based on whatever navigation system is selected for use. The lateral approach mode has an arm-and-capture phase which allows setup of a path intercept with automatic transition into approach capture and track. The vertical guidance is armed for glide-slope capture immediately following localizer capture. Back Course (BC) The back-course localizer approach mode is selected by pressing the BC pushbutton on the AFCS controller. Flight guidance and attitude commands are computed to intercept, capture, and track the back side of the localizer signal. When flying a back-course localizer approach, glideslope capture is automatically inhibited. The back-course mode is set up and flown exactly like a front course with the following differences: • The BC button is selected on the AFCS controller.

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• With the airplane outside the normal localizer capture limits, the advisory display annunciates: • BC in white • HDG SEL in green • At localizer capture, the advisory display annunciates BC in green. The back-course mode is canceled by: • Pressing BC on the AFCS controller • Pressing STBY on the AFCS controller • Selecting APP on the AFCS controller • Coupling to the cross-side HSI • Changing NAV sources

Vertical Modes Air Data Modes (IAS, VS, and ALT) Each of the air data hold modes operates identically. The appropriate mode button is pressed when the desired flight conditions exist. When the hold reference value is established, the system computes commands to maintain this reference value. Only one of these modes can be selected at a time. IAS and VS reference values only can be changed with the pitch wheel while the mode is engaged. Movement of the pitch wheel slews the reference parameter on the advisory display. Depressing the TCS button while an air data mode is engaged allows resynchronization to the value existing when the button is released. Moving the pitch wheel while in the ALT mode causes the message “ALT OFF’’ to be displayed, and the system defaults to pitch hold. Altitude Preselect (ALT SEL) Altitude preselect is enabled by selecting the desired altitude on the preselect controller and depressing the ALT SEL pushbutton. The armed mode is annunciated by a white “ALT

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SEL’’ message on the advisory display. The airplane can be maneuvered toward the selected altitude using pitch, VS, or IAS. The AFCS transitions to the capture mode at a point away from the selected altitude which is dependent upon the vertical speed. The capture mode is annunciated by the green “ALT*’’ message on the advisory display. (The asterisk denotes that the airplane is in the capture portion of the mode operation.) When the airplane has flared onto the new altitude, the system transitions to altitudehold mode, and the asterisk is removed from the message. ALT SEL can be canceled by: • Pressing ALT SEL on the AFCS controller

cleared, and both the GS and LOC are annunciated in green. The glide-slope mode may be canceled by depressing the NAV, APP, or STBY pushbuttons. When the localizer and glide slope both are on track, the radio altitude is less than 1,200 feet, and both NAV receivers are tuned to the same ILS and valid, the AFCS transitions to the dual HSI mode. When this mode is active, both HSI SEL engage arrows are illuminated. In this configuration, both flight guidance channels are using information from both navigation receivers. This allows the approach to be continued in the event of a failure of one NAV receiver. If one receiver fails, the respective engage arrow will extinguish, and the approach will remain active.

• Pressing STBY on the AFCS controller • Selecting go-around • Coupling to the cross-side HSI

NOTE

Navigation Data (V/L, MLS, and AUX) The bottom row of buttons on the AFCS controller are used to select the source of navigation data that is displayed on the HSI and coupled to both flight guidance computers. The NAV SEL button is used to select the leftor right-side systems.

Changing the selected altitude on the altitude alerter while in the capture (*) phase causes the active ALT SEL mode to be canceled and the armed message to be displayed.

Pressing V/L selects the VOR/LOC receiver output to be displayed on the HSI.

Glide Slope (APP) This mode is enabled when the following conditions are met:

Pressing AUX alternately selects RNAV or auxiliary NAV for source data. Modes selected are annunciated on the advisory display.

• A vertical or lateral navigation source is selected. • A localizer frequency is tuned on the NAV receiver. • The APP pushbutton is depressed. • Beam deviation is within the capture envelope. Selection of APP automatically selects the localizer mode. Prior to capture, the armed “GS’’ is annunciated in white and “LOC’’ in green. Upon capture the white message is

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Pressing MLS selects the output of the microwave landing system.

Go-Around (GA) The GA pushbuttons are used to select the goaround mode (see Figure 22-37). Depressing either button disengages the autopilot, and both ADIs display a fixed-pitch command of 10° noseup and wings level or 12° if the flap setting is less than 15° (9° in the series 300 aircraft regardless of flap setting). “GA’’ is annunciated on the advisory display and EADI. The heading select (HDG) mode may be selected while in go-around to replace the wings-level command. Go-around may be canceled by selecting another pitch mode or by depressing the STBY or TCS buttons.

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SAT/TAS DISPLAY

OPERATIONAL MESSAGES

BRT

RESET DISENGAGE/CAUTION/CONDITIONAL STATUS MESSAGES LATERAL ARM MODE STATUS L AFCS

PREVIOUSLY SELECTED LAST SELECTED VERTICAL ARM VERTICAL ARM MODE STATUS MODE STATUS

R AFCS

VERTICAL ACTIVE MODE

LATERAL ACTIVE MODE

Figure 22-39. Advisory Display

SM-300 SERVO DRIVE

TM-400 TRIM SERVO

SM-710 LINEAR ACTUATOR

Figure 22-40. Servoactuators

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Advisory Display The ID-802 advisory display (Figure 22-39) provides the capability for the AFCS to request data from the pilots as well as provide system mode status and annunciation. In addition to SAT/TAS, which is always displayed on the right side of line one, the information displayed includes:

the same unit is used for both the aileron and elevator in the 300 Series aircraft.

SM-710 Linear Actuator The linear actuator translates electrical inputs into a linear mechanical output to control the rudder. The unit provides ram position feedback to the flight guidance computer by means of a linear variable differential transformer.

• AFCS operation and engage status

TM-400 Trim Servo • Flight guidance mode annunciations • System failures and cautions White messages displayed on the left side of line one are displayed for five seconds and are then extinguished. Amber messages displayed on line two are either flashing or steady and require the pilots to acknowledge by clearing the message using the RESET button on the advisory display or in some cases retrimming the airplane or disconnecting the autopilot. Armed FD modes annunciated on line three are in white. Active FD modes annunciated on line four are green. When a mode transitions from the armed state to the captured state, the green message appears in reverse video (black letters on a green background) for a short period of time to emphasize the transition. The brightness is controlled by the BRT knob.

Servoactuators SM-300 Servo Drive The servo drive (Figure 22-40) translates electrical inputs into a clutched rotational mechanical output. This assembly, with a spline output on the clutch, mates with the drum assembly. Tachometer rate signals are fed back to the flight guidance computer servo amplifier. The servo drive unit is bolted to the airplane structure and is connected, through cables, in parallel with the airplane primary control rigging. This unit is used in the elevators and ailerons, but the units are not int e r c h a n g e a b l e ; t h ey h ave d i ff e r e n t d a s h numbers for the 100 Series aircraft; however,

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The trim servo translates electrical inputs into a clutched mechanical rotational output. A spline output on the clutch drives a roller chain. The servo is connected in parallel to the trim surface control rigging by the roller chain drive. The maximum deflection of the trim tab is controlled with limit switches which are part of the trim servo.

AFCS Failures The AFCS includes monitors which assess the validity of the computer and other system components. The monitors are in addition to those which provide fail-passive and failoperational operation. The power supply, microprocessors, and other hardware within the system are monitored for availability. If a failure is detected, the system is rendered inoperative. With application of power to the AFCS, it goes through a power-up preflight test; the message “SYSTEM TEST’’ is displayed on the advisory display. Upon successful completion of the test by both computers, the message “L AFCS MASTER’’ is displayed. If a failure is detected in either flight guidance computer, the respective “L’’ or “R AP/YD FAIL’’ message is displayed. A failure of both computers results in both FAIL messages or a series of dashes being displayed on the advisory display. A failure within the flight guidance computer which prevents operation of the flight director function results in an FD flag on the ADI or EADI.

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Figure 22-41. NAV Control Head

DME

INN OUT

HOLD

MID

Figure 22-43. MARKER SENS Panel

TAS

REVN

Figure 22-42. Marker Beacon Lights

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If a failure occurs in a flight guidance computer after the initial power-up test, any attempt to engage the autopilot results in the “L’’ or “R AP/YD FAIL’’ message. Under this circumstance the yaw damper may still be operational, depending upon the nature of the failure. If depressing the YD button displays the “L’’ or “R AP/YD FAIL’’ message, the computer failure affects the yaw damper function.

NAVIGATION EQUIPMENT GENERAL Navigation equipment may include the flight management computer (FMC) and dual combined VOR/marker beacon receivers, dual DME receivers, and dual ADF receivers. Miscella-neous equipment covered in this section includes the transponder, cockpit voice recorder, flight data recorder, ground proximity warning system, and emergency locator transmitter.

NAVIGATION SYSTEM RECEIVERS General Navigation receivers include dual combined VOR/marker beacon receivers, dual DME receivers, and dual ADF receivers. The marker beacon, ADF, VOR, and RNAV are controlled by individual control panels on the center console and the glareshield. The dual DMEs are selected simultaneously by respective VOR receiver selection. When electromechanical instruments are installed, VOR/DME 1 deviation and distance a r e d i s p l a y e d o n t h e p i l o t ’s H S I , a n d VOR/DME 2 deviation and distance are displayed on the copilot’s HSI. VOR and ADF bearing information is displayed on the RMIs. Marker beacon receiver reception is annunciated by marker beacon lights above each ADI. On EFIS-equipped aircraft, either VOR/DME

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1 or VOR/DME 2 information can be displayed on either EHSI. As well, ADF 1 and ADF 2 can be displayed on either EHSI, and the EADI displays marker beacon lights.

VOR Navigation System D u a l VO R n a v i g a t i o n s y s t e m s p r o v i d e reception of VOR, localizer, glide-slope, and marker-beacon signals. The VORs are controlled by two NAV control heads located in the glareshield. Each control head (Figure 22-41) p r ov i d e s a c t ive f r e q u e n cy s e l e c t i o n a n d standby frequency selection. Normally, a frequency is selected in the SBY display and then transferred to the active display for navigation receiver operation by pressing the transfer pushbutton (the button with the two-headed arrow). When the transfer pushbutton is momentarily pressed, the frequency in the active display and the frequency in the standby display exchange places. The larger knob of the concentric frequency selector knobs increases or decreases the MHz portion of the frequency in one-MHz increments. The MHz and kHz portion of each display rolls over to the opposite end of the frequency band when selected beyond the upper or lower limits of the display. The kHz display rolls over from .00 to .95 or .95 to .00, and the MHz display rolls over from 108 to 117 or 117 to 108.

Marker Beacon System The dual marker beacon receivers are integral with the dual VOR/localizer navigation receivers. The marker beacon receivers illuminate a blue legend (OUT) at the outer marker, an amber legend (MID)at the middle marker, and a white legend (INN) at the inner/airway marker. The marker beacon lights (Figure 2242) are located directly above each ADI. Marker beacon sensitivity may be selected to HI or LO by the NO. 1 and NO. 2 switches on the MARKER SENS panel (Figure 22-43) on the center console.

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Figure 22-44. ADF Control Head

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Distance Measuring Equipment (DME)

NOTES

The dual DMEs provide distance information from the selected VOR navigation stations. The DME transceivers are controlled by the NAV control heads on the glareshield. When a VOR frequency is selected, a paired DME frequency is selected simultaneously. This gives slant range distance to the VOR on the HSI(s) when VOR is selected to that HSI. DME frequency can be held by each DME transceiver while the associated NAV controller is reselected to another frequency by pressing the pilot’s or copilot’s DME HOLD switchlight above each ADI (see Figure 22-42).

Automatic Direction Finding System (ADF) The dual ADF systems allow automatic tothe-station relative bearing and radio broadcast reception. The ADF systems are controlled by the two ADF control heads (Figure 22-44) located on the center console. Each control head allows mode selection and display, frequency selection and display, and channel programming (optional).

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22-64 8 6 7 10

9 5

GND CREW

GPWS

FWD AFT 18

17

11 13 19

15 16 14 12

1 VHF1

BLKHD 182.0

1-

HF

VHF2

VOR -2 MLS

UHF

INT

VHF1

FM

SERV/INT

1- DME

-2

1-

-2

MICROPHONES BOOM MASK

MICROPHONES BOOM MASK

HEADPHONES AUX NORMAL

HEADPHONES AUX NORMAL

PA MIC SPKR

VOL

WARDROBE

MKR BOOM HOT MIC

MASK

AFD

ON

20

2

3

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

4

GROUND CREW JACK

1. GROUND CREW JACK, AFT (NACELLE) 2. GROUND CREW JACK, FWD (NOSE) 3. REMOTE ELECTRONICS UNIT 4. GROUND CREW ADVISORY LIGHT 5. PILOT'S PRESS-TO-TALK SWITCH (NOSEWHEEL STEERING CONTROL) 6. PILOT'S PRESS-TO-TALK SWITCH (CONTROL COLUMN) 7. COPILOT'S PRESS-TO-TALK SWITCH (CONTROL COLUMN) 8. COPILOT'S PRESS-TO-TALK SWITCH (COPILOT'S SIDE PANEL) 9. PILOT'S COCKPIT SPEAKER 10. COPILOT'S COCKPIT SPEAKER 11. PILOT'S HAND MICROPHONE 12. COPILOT'S HAND MICROPHONE 13. PILOT'S MICROPHONE JACK 14. COPILOT'S MICROPHONE JACK 15. PILOT'S AUDIO CONTROL PANEL 16. COPILOT'S AUDIO CONTROL PANEL 17. OBSERVER'S AUDIO CONTROL PANEL 18. PILOT'S JACK PANEL 19. COPILOT'S JACK PANEL 20. OBSERVER'S JACK PANEL

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Figure 22-45. Audio Integrating System

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COMMUNICATIONS EQUIPMENT AUDIO INTEGRATION General The audio integrating system provides the pilot, copilot, and observer with the means to control the airplane communication radio systems for transmission and reception and to monitor the airplane navigation receivers. In addition, interphone communication is provided among the crew and between the crew and ground crew. The audio control panel is the principal component of each station. Figure 2245 shows the components of the audio integrating system.

Description and Operation Each pilot station has the following audio controls: • Audio control panel (Figure 22-46) • Hand microphone • Jack panel • Cockpit speaker • Press-to-talk interphone switch An additional press-to-talk switch, labeled “PTT,’’ is provided for the pilot on the nosewheel steering control, and an interphone-transmit switch, labeled “INPH/XMIT,’’ is provided for the copilot on the copilot’s side panel. The observer’s station has the following audio controls:

• Audio control panel • Jack panel • Transmit-interphone toggle switch • Microphone jack • Headset jack

The audio system also contains a ground crew jack located on the forward left fuselage and an additional ground crew jack located in the refuel-defuel panel in the right engine nacelle. Common to all stations and jacks is the primary interface component of the audio integrating system, called the “remote electronics unit,’’ located in the avionics rack. Each audio control panel contains a rotary MIC (microphone) transmission selector switch which provides the selection of VHF 1, VHF 2, INT (interphone), and optional HF, UHF, and FM communication systems. The panel also contains 14 push-on, push-off, turn (volume control) receiver selector switches and two additional switches of the same type for speaker volume and service interphone. The panel has a leverlock microphone selector switch which provides for the selection of BOOM or MASK and a switch of the same type for the selection of HOT MIC operation. Operation of the controls and switches on the audio control panels control the functioning of the remote electronics unit to which all the communication and navigation system audios are connected and through which all audio signals are routed.

SERVICE INTERPHONE The service interphone is for on-ground communication and is operated through the audio control panel. The system has two external ground crew stations (left side forward fuselage and refueling-defueling panel in the right nacelle). Volume control is possible from t h e p i l o t ’s a n d c o p i l o t ’s s t a t i o n s o n l y. Interphone communication from any of these three stations requires that the operator select INT on the rotary MIC switch and press XMIT, PTT, the hand microphone switch, or INPH. If INT is not selected on the rotary MIC selector switch, then interphone is possible only by pressing INPH. Either action permits all flight crew stations to receive interphone audio by pressing the SERV/INT selector switch.

• On-off switch (observer’s light)

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PUSH ON/PUSH OFF TURN VOL SPEAKER SELECTOR SWITCH AND VOLUME CONTROL

VHF1

1-

PUSH ON/PUSH OFF TURN VOL SERVICE/ INTERPHONE SELECTOR SWITCH AND VOLUME CONTROL

MICROPHONE/INTERPHONE SELECTOR SWITCH

HF

VHF2

VOR -2 MLS

UHF

INT

VHF1

FM

SERV/INT

1- DME

-2

1-

-2

PA MIC SPKR

MKR BOOM

VOL

BOOM/MASK MICROPHONE SELECTOR SWITCH

HOT MIC

MASK

ADF

ON

HOT MIC SELECTOR SWITCH (100 SERIES SWITCH LABELED NORM/EMER)

PUSH ON/PUSH OFF TURN VOL RECEIVER SELECTOR SWITCH AND VOLUME CONTROLS COMBINED (TYPICAL 14 PLACES)

SERIES 100A/300

Figure 22-46. Audio Control Panel (Series 100A/300)

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The ground crew forward station and the ground crew station on the refueling-defueling panel receive interphone audio at all times. A ground crew advisory light (at the left side of the nosewheel steering arm) comes on and indicates FWD (forward) when a ground crew station headset/microphone is plugged into the jack labeled “GROUND CREW JACK’’ on the left side of the aircraft nose. A light labeled “A F T ’’ c o m e s o n w h e n a h e a d s e t / m i c i s plugged into the jack on the refueling-defueling panel. The FWD and AFT advisory lights alert the flight crew that interphone communication with a ground crew station may be made and also indicates which station. If either the pilot’s or the copilot’s audio control panel fails, the affected crewmember can monitor the audio integrating system by plugging his headset into the HEADPHONES-AUX jack on his jack panel.

VHF COMM RADIOS Air-to-ground communications are achieved using dual very high frequency transceivers (VHF COMM). The two control heads are located in the center console. The pilot, copilot, and observer can transmit and receive on the system through their audio control panels, using headsets and boom microphones. The crew can also use the overhead loudspeakers and hand-held microphones. The COMM system provides voice communications in 25-kHz increments. The control h e a d d i s p l a y s a c t ive a n d p r e s e t s t a n d b y frequencies. Normally a frequency is selected in the SBY display and then transferred to the active display by pressing the transfer pushbutton (Figure 22-47). Programmable channels are optional. To transmit on the radio, rotate the MIC selector switch to the desired transmitter. Microphone selection is made by setting the MIC toggle switch to either MASK or BOOM. With the desired radio frequency selected at the appropriate audio control panel, transmission by the pilot is possible by holding the

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PTT/INPH switch (control wheel) at PTT, pressing and holding the PTT switch on the nosewheel steering arm, or using the pilot’s hand microphone. Operation of the pilot’s hand microphone overrides the pilot’s boom and mask microphones. Transmission by the copil o t i s p o s s i b l e b y h o l d i n g t h e c o p i l o t ’s PTT/INPH switch (control wheel) at PTT, setting and holding the INPH/XMIT switch (copilot’s side panel) to XMIT, or using the copilot’s hand microphone. Operation of the copilot’s hand microphone overrides the copilot’s boom and mask microphones. During transmission, sidetone from the transmitter is applied to the audio integrating system. In addition, the pilot’s and copilot’s microphone audio is constantly applied to the cockpit voice recorder, regardless of the pos i t i o n o f e i t h e r t h e p i l o t ’s o r c o p i l o t ’s PTT/INPH selection. For communications radio reception, the desired radio receiver audio output is selected by pressing in the appropriate pushon, push-off, turn (volume control) switch. Cockpit speaker audio is selected by pressing in the SPKR/VOL switch.

NOTE Speaker audio is muted when the respective PTT/INPH switch on either control wheel is selected to PTT or INPH or when the PTT switch on the nosewheel steering switch is pressed.

NOTE With cockpit speakers selected, care must be taken when adjusting the volume control because acoustic feedback may cause a loud squeal to be heard if it is set too high.

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Figure 22-47. VHF COMM Control Head

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VHF NAV RADIOS

NOTES

For navigation radio reception, the desired navigation radio receiver audio output is selected by pressing in the appropriate push-on, push-off, turn (volume control) switch. Navigation radio audio output from the remote electronics unit is applied to the cockpit voice recorder and headphones or cockpit speaker, as desired.

EMERGENCY OPERATION Emergency operation is selected automatically. Emergency operation is automatically selected if a power failure occurs within or to the remote electronics unit. In emergency operation, the pilot’s headphone is connected directly to VHF 1 and VOR 1/MLS 1 receivers and the pilot’s microphone is connected directly to the VHF 1 microphone input so that both audio signals bypass all amplifiers or other active circuitry normally in the audio path. The copilot’s and observer’s headphones are connected directly to VHF 2 and VOR 2/MLS 2 receivers, and the microphones are connected directly to the VHF transmitter. The pilots’ hand microphones are not operational during emergency operation. Either the boom or mask microphone must be used for emergency operation and the unused (boom or mask) microphone disconnected at the phone jack. The interphone system is not operational during emergency operation.

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1 9 CABIN INTERPHONE

11

CHIME

PA

CALL

EMER

6 15

17

5

EMER

10 16

PA

8 CALL

12

B O A R D ' g

13 ATTENDANT'S CALL SWITCH

L I G H T S

OVERHEAD STOWAGE BIN ATTENDANT'S CALL LIGHT 5 3 4

BLKHD 182.0

VOL P A

2

SENS

14 WARDROBE

7

RETURN TO SEAT SIGN

VIEW LOOKING FWD LAVATORY

1. FLIGHT CREW CABIN INTERPHONE CONTROL UNIT 2. PASSENGER ADDRESS AMPLIFIER 3. VOLUME CONTROL (PA AMPLIFIER) 4. SENSITIVITY CONTROL (PA AMPLIFIER) 5. LAVATORY SPEAKER 6. CABIN SPEAKERS (FOUR PLACES) 7. REMOTE ELECTRONICS UNIT 8. FLIGHT CREW HANDSET 9. FLIGHT ATTENDANT'S HANDSET 10. HANDSET CRADLE 11. HANDSET CRADLE 12. PASSENGER SERVICE UNIT (NINE PLACES, LEFT SIDE) 13. PASSENGER SERVICE UNIT (NINE PLACES, RIGHT SIDE) 14. SERVICE CALL PUSHBUTTON 15. FLIGHT ATTENDANT'S CONTROL PANEL 16. FLIGHT ATTENDANT'S ANNUNCIATOR PANEL 17. SMOKE DETECTOR LIGHT

Figure 22-48. Passenger Address Components and Controls (Sheet 1 of 3) SMOKE DETECTOR LIGHT

EMER

CABIN INTERPHONE PA

CHIME

CALL

CALL/EMER

PA

PA EMER

SERVICE CALL (LAVATORY)

CALL

CABIN INTERPHONE CONTROL UNIT

SERIES 100

B O A R D ' g

L I G H T S

FLIGHT ATTENDANT'S CONTROL PANEL

PSU CALL

FLIGHT ATTENDANT'S ANNUNCIATOR PANEL

Figure 22-48. Passenger Address Components and Controls (Sheet 2 of 3)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PASSENGER ADDRESS SYSTEM

NOTES

The passenger address system is operated from the cockpit passenger address and cabin interphone panel. The system draws the attention of the pilot by a visual and audible call. The passenger address system provides the following: • Flight crew to passengers address • Flight attendant to passengers address • Passenger to flight attendant call • No smoking and fasten seat belts warning chime In addition, provision is made for passenger announcement and boarding music tape player inputs. Figure 22-48 illustrates the passenger address components and controls.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

EMER

EMER

FLIGHT CREW CABIN INTERPHONE CONTROL PANEL

PA

PA

CALL

ATT

CALL

EMER

ATTENDANT'S NO. 2 CONTROL PANEL

ATT

B O A R D I N G

PA

CALL

ATT

L I G H T S

ATTENDANT NO. 2 ANNUNCIATOR PANEL (CEILING CENTERLINE AFT OF BULKHEAD) PINK

RED

GREEN

GREEN

GREEN AMBER PINK

RED

ATTENDANT'S NO. 1 CONTROL PANEL

FOR SYSTEM ANNUNCIATION, SEE ATTENDANT NO. 2 ANNUNCIATOR PANEL

BLUE GREEN

ATTENDANT NO. 1 ANNUNCIATOR PANEL

AMBER

LEGEND

BLUE

LAVATORY SMOKE (FLASHING) OXYGEN PRESSURIZED DROP DOWN OXYGEN ONLY EMERGENCY CALL PA ACTIVE LAVATORY CALL MASTER CALL

SERIES 300

Figure 22-48. Passenger Address Components and Controls (Sheet 3 of 3)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AUDIO INTEGRATING, PASSENGER ADDRESS, AND ATTENDANT’S INTERPHONE— SERIES 300

NOTES

The audio control panels in Series 300 and Series 100A aircraft no longer have a NORM–EMER switch. This has been replaced with a HOT MIC–ON switch. This switch allows the pilot, copilot, and observer to continuously key their microphone while keeping their hands free in the interphone mode only (see Figure 22-46). Also, the audio control panel has another microphone selection marked “PA.’’ Selecting the microphone selector to PA allows use of the boom or mask microphone to make PA announcements. There is no PA handset in the cockpit of the Series 300 aircraft. The flight crew cabin interphone control unit is also changed. The four switchlights are now labeled “ATT,’’ “CALL,’’ “PA,’’ and “EMER.’’ The pilot now pushes PA and then CALL to make a private call to either attendant without it being broadcast over the PA system. The first attendant’s control panel has an additional switchlight marked “ATT,’’ which is used to call from one attendant to the other. A second attendant’s control panel annunciator panel, and handset are in the rear of the aircraft (aft of divider bulkhead X578.000).

Revision 2

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Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-49. Static Discharge Wicks

Figure 22-50. XPDR Control Head

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

STATIC DISCHARGE WICKS

NOTES

A static electrical charge, commonly referred to as “P’’ (precipitation static), builds up on the surface of an airplane while in flight and causes interference in radio and avionics equipment operation. The charge may be dangerous to persons disembarking after landing as well as to persons performing maintenance on the airplane. The static wicks are installed on all trailing edges (Figure 22-49) and dissipate the static electricity in flight. Two types of static dischargers are used on the Dash 8: one is approximately six inches in length, and the other is eight inches long. Each discharger consists of a resistive-coated fiberglass rod with a pellet of composite carbonbased material at the tip. A total of 24 static discharge wicks are installed on the Dash 8.

MISCELLANEOUS EQUIPMENT TRANSPONDER The single (dual, optional) transponder responds to air traffic control (ATC) facility interrogations with a selected ATC-assigned code to show controllers the present position of the airplane. Altitude can also be shown with mode C. The receiver/transmitter is controlled by the XPDR control head (Figure 22-50) located in the center pedestal. The transponder control head provides operating mode selection, ATC code selection, mode and code display, and an ident button. The mode selector knob on the control head manually selects the standby (SBY), ON, altitude reporting (ALT), and test (TST) modes.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-51. CVR Monitor Panel

ULD AUDIO INTEGRATING SYSTEM PASSENGER ADDRESS AND ENTERTAINMENT SYSTEM 18 VDC (PRE-AMP POWER)

MIKE

PILOT'S AUDIO (CHANNEL 3) COPILOT'S AUDIO (CHANNEL 2) PA AUDIO (CHANNEL 1)

COCKPIT MIC AUDIO (CHANNEL 4)

COCKPIT VOICE RECORDER CONTROL UNIT

RECORDER AUDIO MONITOR TEST METER SIGNAL

COCKPIT VOICE RECORDER

ERASE COMMAND

ERASE INTERLOCK PARK BRAKE

(EQ50)

NC GND C NO NC C S2 NO CENTER CONSOLE

PROXIMITY SWITCH ELECTRONIC UNIT (PSEU)

3351-S1

AVIONICS CIRCUIT BREAKER PANEL CVR 28 VDC (A3) LEFT MAIN BUS CVR (A10) 115 VAC RIGHT BUS

P/O OVER TEMP PCB-B K7

28 VDC (RELAY POWER) 115-V, 400-HZ PRIMARY POWER

NOTES: 1. PARK BRAKE SWITCH 3351-S1 IS SHOWN IN BRAKES– ON POSITION. 2. COMPONENTS NOT KEYED NUMERICALLY ARE IDENDIFIED IN APPROPRIATE SECTIONS OF MAINTENANCE MANUAL. 3. UNDERWATER LOCATING DEVICE (ULD) REQUIRES NO EXTERNAL CONNECTIONS. 4. THE LOOP CIRCUIT IN PROXIMITY SWITCH ELECTRONIC UNIT CLOSES WHEN CONDITIONS IN EQUATION 50 ARE SATISFIED. EQUATION 50: (LGWOW 1 + LGWOW 2) (RGWOW 1 + RGWOW 2) (NGWOW 1+ NGWOW 2) SYMBOLS: = AND, + = OR FOR FULL DETAILS OF PSEU LOGIC AND EQUATIONS, REFER TO CHAPTER 32 IN THE MAINTENANCE MANUAL.

COCKPIT VOICE ULD RECORDER (CVR)

MONITOR JACK

K4

GND ACCELERATION SWITCH (SYSTEM CODE 3401) AUTOFLIGHT CONTROL SYSTEM INTERFACE UNIT COCKPIT VOICE RECORDER CONTROL UNIT

TEST ERASE PUSH-BUTTON PUSH-BUTTON SWITCH SWITCH METER

3

MICROPHONE

TEST ERASE

HEADSET

HEADSET JACK

Figure 22-52. Cockpit Voice Recorder System

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

COCKPIT VOICE RECORDER

NOTES

The cockpit voice recorder (CVR) microphone and controls are located in the center pedestal (Figure 22-51). The recorder unit is located in the tail in a crash-survivable case adjacent to the flight data recorder. An inertia switch is located on the front spar above the cabin fuselage. The system operates any time the airplane has power on it and records a continuous loop 30 minutes in length. Playback is not possible u n l e s s t h e r e c o r d e r i s r e m ove d f r o m t h e airplane. Recording is automatically stopped by the inertia switch in the event the switch is actuated. An underwater beacon is attached to aid in locating the CVR. The system can be tested any time it is operating by pressing the TEST button while the headphones are plugged into the recorder headphones jack. When the button is pressed, two 400-Hz test tones are heard in the headphones. At the same time, the meter pointer deflects to full scale twice. The tape can be erased at the end of a flight by pressing the ERASE button (optional) for at least 16 seconds after the airplane is on the ground and the aircraft parking brake is set. Figure 22-52 shows a block diagram of the system.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT DATA RECORDER

ELT. OFF AUTO

NORM

ON

GND TEST

WARNING EMERGENCY USE ONLY UNAUTHORIZED OPERATION PROHIBITED

Figure 22-53. FDR/ELT Control Panel

12B 12A 12H

12G 12E 6

2

12C

3 18 4

1

7

8,9,10,11

12D 12F

*5

19 13 17 12

NO. 1 2 3 4 *5 6 7 8 9 10 11

14

15,16 CHAPTER DESCRIPTION REF FDAU 31 3-AXIS ACCELEROMETER 31 DFDR 31 DFDR TEST SWITCH 31 31 FLIGHT DATA ENTRY PANEL INERTIA SWITCH 31 COPY RECORDER CONNECTOR 31 PROP RPM NO. 1 AND NO. 2 61 ENGINE TORQUE NO. 1 & NO. 2 77 77 GAS GENERATOR RPM NO. 1 AND NO. 2 27 FLAP POSITION INDICATOR

* WHEN INSTALLED FROM ENGINE INSTRUMENT PANEL CHAPTER NO. DESCRIPTION REF 12 13 14 15 16 17 18 19

PWRD FLT CONT SURF INDICATORS PROXIMITY SW ELECTRONICS UNIT FLIGHT GUIDANCE COMPUTER VHF COMM NO. 1 VHF COMM NO. 2 ELEVATOR TRIM WHEEL POTENTIOMETER ANTICOLLISION LIGHTS SWITCH AFCS INTERFACE UNIT

27 32 34 23 23 27,31 33 34

Figure 22-54. Flight Data Recorder System

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT DATA RECORDER The digital flight data recorder assesses, measures, and records parameters of flight for subsequent analysis or investigation of incidents or accidents. The recorder is in a crashsurvivable case in the tail. The system is powered on the ground when there is power on the airplane and the anticollision lights are on; after the airplane is airborne, anticollision light operation is not required for recorder operation. Controls for operation are located on the overhead panel and are shown in Figure 22-53. Component locations and a block diagram are shown in Figure 22-54 and 22-55.

DFDR TEST PANEL

6

NORM TEST

ANTI-COLLISION LIGHTS OFF 0N 18

FDR STAT 28V

115V 400

The recorder records 25 hours of information in a continuous loop. Playback is accomplished by maintenance personnel only. The flight data recorder has an underwater locator beacon which emits a sonar “ping’’ when immersed in water. The system can be checked for serviceability without the anticollision lights on by momentarily pressing the GND TEST switch on the FLIGHT DATA RECORDER panel on the overhead panel. During the test, the FLT DATA RECORDER caution light goes out, indicating that the system is operational. An inertia switch will deenergize the system in the event of a crash.

FDR

19 115 V, 400 HZ IN DFDR FAIL

13 WOW

P/O PSEU INERTIA SWITCH EQUATION 42 EQUATION 43 DATE AND FLIGHT NUMBER ACARS DATA

GMT (OPTIONAL)

FOR DETAILS OF INPUTS REFER TO CHAPTERS 32 AND 34 OF THE MAINTENANCE MANUAL.

GMT TIME SIGNAL

FLT DATA 28 V RECORDER

3

PLAYBACK DATA 5 VDC, +12 VDC, –12 VDC 115 V, 400 HZ DIGITAL FLIGHT DATA RECORD INHIBIT RECORDER 1 TRACK DATA OUT DIRECTION PLAYBACK HIGH SPEED SELECT MAINT FLAG ODD/EVEN DIRECTION

6

ACARS (OPTIONAL)

LOGIC

P/0 AFCS INTERFACE UNIT

EQUATION 79

*5 FLIGHT DATA ENTRY PANEL (OPTIONAL)

P/O CAUTION LIGHTS PANEL

COPY RECORDER CONNECTOR 3133-J7 (OPTIONAL)

TRACK SELECT P/O FLIGHT DATA ACQUISITION UNIT

AIRPLANE SWITCHES (PSEU) RADIO ALTIMETER

AFCS SYSTEM

P/N MS24264R16B24SN

"A" DATA BUS

TCS AND AUTOPILOT CONTROL "B" DATA BUS

Figure 22-55. Flight Data Recorder in Block Diagram

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22-80 AUDIO OUT (LO LEVEL)

NO. 1 VHF NAV CONTROLLER

1 AUDIO OUT (HI LEVEL)

VARIABLE 3 ATTENUATOR

GPW VISUAL ANNUNCIATE

GS MODE ENABLE GROUND

TO AUDIO INTEGRATING SYSTEM

LH SIDE

SELF TEST

LH SIDE

GS ADVISORY

LH SIDE

4 SPEAKER

PULL UP

AFCS CONTROL PANEL

GS VALID

2 GPWS FLAP O/RIDE SWITCH

RADIO ALT RX/TX NO. 1 PILOT ADI

GROUND GS CANCEL PROXIMITY WARNING COMPUTER AFCS INTERFACE GPWS UNIT FAIL K5

LH SIDE

5

3 (MOUNTED ON REAR OF PANEL)

LIGHT EMG EXIT

GPWS TEST

BELOW G/S

4

LIFE VEST

FWD

FLIGHT COMPARTMENT THRESHOLD CEILING PANEL (LOOKING UP AT OVERHEAD PANEL)

6

BLKHD 182.00

RH SIDE SIMILAR

P/O MASTER CAUTION PANEL 1

GPWS STALL WARNING HORN

FLAP POSITION >15° GEAR POSITION (DOWN & LOCKED)

WARDROBE

STALL PILOT'S FWD SIDE CONSOLE WARNING (LOOKING FWD) COMPUTERS NO. 1 & NO. 2 INDEX NO.

AFCS RADIO ALTITUDE VALID INTERFACE RAD ALT VALID UNIT RADIO ALTITUDE RADIO ALTITUDE RAD ALT TEST RAD ALT TEST RAD ALT TEST DECISION HEIGHT GND

GLARSHIELD PANEL (LOOKING FWD) 6 CL

X180.9–

GS MODE INHIBIT

PROXIMITY SWITCH ELECTRONICS UNIT (PSEU)

6

X195.7–

GS DEVIATION NO. 1 VHF NAV RECEIVER

5

RAD ALT TEST

CO-PILOT ADI

1 2 3 4 5 6

PART NO.

MANUFACTURER

965-0476-088 SUNDSTRAND 99-522-K83-15099 AEROSPACE OPTICS OHMITE CLU5001 OAKTRON IND T4053 851-27864-001 MASTER SPECIALTIES 851-27864-002 MASTER SPECIALTIES

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AIRSPEED ALTITUDE RATE VALID ALTITUDE RATE VARIABLE ALTITUDE RATE REFERENCE

NO. 1 AIR DATA COMPUTER

5

Canada Ltd.

Revision 4

Figure 22-56. Ground Proximity Warning System

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

GROUND PROXIMITY WARNING SYSTEM General The ground proximity warning system (GPWS) provides visual, aural, and synthesized voice annunciation to warn of an impending hazardous situation with regard to terrain avoidance. It is powered any time there is power on the airplane.

The GPWS computer receives inputs from the radio altimeter, No. 1 ADC, No. 1 glideslope receiver, and landing gear and flap position switches. The inputs are continuously processed to provide airplane flight path surveillance when between 50 and 2,450 feet AGL. With the exception of glide-slope Mode-5 violations and Mode 4B “too low flaps,’’ all warnings can be canceled only by taking corrective action such as adding climb power and executing a positive pull-up. Warning of system failure is provided by a GPWS caution light.

Operation There are six modes of operation: • Mode 1 is conditioned for excessive rates of descent with respect to terrain. The pull-up warning annunciator light illuminates, and the voice advisory is “sink rate, sink rate,’’ followed by “whoop, whoop, pull up.’’ • Mode 2 is conditioned for excessive closure rates to terrain. The pull-up warning annunciator light illuminates, and t h e vo i c e a d v i s o r y i s “ t e r r a i n , t e r rain–whoop, whoop, pull up, terrain.’’

Revision 2

• Mode 3 is conditioned for significant altitude loss before acquiring a predetermined terrain clearance after takeoff or missed approach. The pull-up warning annunciator light illuminates, and the voice advisory is “don’t sink.’’ • Mode 4 is conditioned for insufficient terrain clearance based on airplane configuration and speed (inadvertent proximity to terrain). Mode 4 is further divided into Modes 4A and 4B. • Mode 4A is conditioned for inadvertent proximity to terrain with the landing gear up. The pull-up warning annunciator light illuminates, and the voice advisory is “too low terrain’’ or “too low gear,’’ depending on speed and altitude. • Mode 4B is conditioned for inadvertent proximity to terrain with flaps up and landing gear down. The pull-up warning annunciator light illuminates, and the voice advisory is “too low flap’’ or “too low terrain,’’ depending on speed and altitude. For zero-flap landings, a flap override switch is provided adjacent to the nosewheel steering handle on the pilot’s left console. • Mode 5 is conditioned for inadvertent descent below glide slope on an ILS approach. The below-glide-slope annunciator light illuminates, and the voice advisory is “glide slope.’’ The warning can be canceled by pressing the glides l o p e i n h i b i t s w i t c h , bu t i t w i l l b e rearmed if the airplane recaptures the glide-slope beam. • Mode 6 is conditioned for descent below minimum selected altitude. The voice advisory is “minimums, minimums,’’ and the decision height amber advisory lights are illuminated.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

REMOTE CONTROL CABLE

ANTENNA

FWD Z210.54 EMERGENCY LOCATOR TRANSMITTER

ANTENNA COAXIAL CABLE

REAR BRACKET CLIP WITH QUICK RELEAS Z198.6

REMOTE CONTROL CABLE

X681.0

X698.55

FLIGHT DATA RECORDER

E.L.T. OFF

NORM

AUTO GND TEST

WARNING EMERGENCY USE ONLY ON UNAUTHORIZED OPERATION PROHIBITED

Figure 22-57. Emergency Locator Transmitter

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Due to the possibility of activating more than one warning condition at a time, there is a priority in the voice advisory messages: 1. “Whoop, whoop, pull up’’—Modes 1 and 2

illuminates, the advisory voice “glide slope’’ occurs once followed by a onesecond pause, the pull-up light comes on, the advisory voice “whoop, whoop, pull up’’ continues for several repetitions and then stops, and the pull-up light then goes out.

2. “Terrain’’—Mode 2 3. “Too low terrain’’—Mode 4

6. Release the test switch, and the belowglide-slope light and GPWS monitor light go out.

4. “Too low gear’’—Mode 4 5. “Too low flap’’—Mode 4 6. “Minimums, minimums’’—Mode 6 7. “Sink rate’’—Mode 1 8. “Don’t sink’’—Mode 3 9. “Glide slope’’—Mode 5 Figure 22-56 shows a block diagram of the ground proximity warning system.

Troubleshooting To perform a GPWS self-test: 1. Ensure that the GPWS circuit breaker is in and that the radio altimeter and No. 1 ADC are installed and operational. 2. Ensure that the landing gear is not retracted and that the flaps are up. 3. Ensure that the GPWS flap override switch is closed (not in override). 4. Ensure that the ground proximity monitor light is off. If it is on, press and release the test switch to cycle the test sequence. If it remains on, perform troubleshooting per the Maintenance Manual. 5. Press and hold the pull-up/GPWS test switchlight. The GPWS monitor light illuminates, the below-glide-slope light

Revision 2

Any deviation from the test may not mean a ground proximity warning computer failure. Sensor faults and/or broken/missing wire connections can cause erroneous test sequencing.

EMERGENCY LOCATOR TRANSMITTER The emergency locator transmitter (ELT) is a self-contained unit capable of manual or auto matic operation. Automatic operation is caused by an inertia switch. Manual operation is from the ELT control panel (see Figure 22-57) on the overhead panel. If necessary, the ELT can be removed from the dorsal fin and operated with a switch on the set. During the prestart check, the ELT switch should be selected to AUTO and 121.5 MHz (guard) monitored to ensure that there is no ELT signal. For portable operation after the set has been r e m ove d f r o m t h e a i r p l a n e , t h e p o r t a b l e antenna must be attached to the transmitter and the switch on the unit turned on. An indicator light on the unit comes on to indicate that the ELT is transmitting.

NOTE The off position of the remote switch is operative only when power is being applied to the airplane DC bus. Figure 22-57 illustrates the emergency locator transmitter system.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

TARGET ALERT INDICATOR ON/OFF BRIGHTNESS

STANDBY/ TEST WEATHER/ CYCLIC

AZIMUTH MARKERS

RANGE MARKERS

VARIABLE GAIN "ON" INDICATOR

BRT

T

OFF SB/T

VAR

WX/C

20

MAP FUNCTION

MAP

RAIN ECHO COMPENSATION

RCT

GROUND CLUTTER REDUCTION

GCR

200

100 50

15 +7.3

WX

VAR

SENSITIVITY CONTROL

MODE FIELD

VARIABLE GAIN ON/OFF

TGT ALT

TARGET ALERT ON/OFF

AZ MK

AZIMUTH MARKERS ON/OFF

RANGE SELECTOR BUTTONS

25

10 5

GAIN

COLOR BAR

300

25

10 TILT

SCAN

SCAN SELECT 60/120°

ANTENNA TILT INDICATOR

ANTENNA TILT CONTROL

Figure 22-58. Weather Radar Indicator

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

WEATHER RADAR General The Primus 800 weather radar system includes the receiver-transmitter, the digital indicator, and the antenna unit. The Dash 8 weather radar is an X-band radar designed for weather detection and analysis. The indicator (Figure 22-58) displays storm intensity levels in bright colors contrasted against a black background. Areas of heaviest rainfall appear in red, medium intensity appears in yellow, and weakest intensity returns are green.

Ground mapping is possible with the radar. In MAP mode, prominent landmarks are displayed which may allow identification of landwater contrast, mountainous areas, and large population centers. Video levels of increasing reflectivity are displayed as black, cyan, yellow, and magenta. TILT and GAIN must be carefully adjusted and balanced for different terrain types to get the optimum picture. Radar range is 10 to 300 nautical miles, but consistent, optimum weather returns are painted best at shorter ranges. When the radar is operated in conjunction with the optional EFIS, radar returns may be displayed on the EHSI.

Troubleshooting CAUTION Observe safety precautions when the radar is on. A radiation hazard exists in all modes except standby and off.

1. Ensure that the radar circuit breakers are in and that the AHRS is operational.

Revision 2

2. Ensure that all personnel are aware that the radar will be tested and that a radiation hazard exists. Ensure that the area in front of the airplane is clear of all flammable material. Do not allow any unqualified personnel in the immediate area without supervision. The area which must be clear is 6.5 feet for 360° and 100 feet in the sector.

Fault Displays Operation of the radar system is monitored by fault circuits which warn if an abnormal condition exists. Visual fault displays are provided for four types of monitored faults: display colors, display sensitivity, transmitter and receiver integrity, and antenna synchronization. Display colors are continuously monitored by the color bars. If a fault causes a change to unfamiliar colors, the severity level denoted by the faulty color(s) is directly coded by the bar. Video detection levels for the display are monitored by the target bands in the test pattern. For example, if the detector for severity-level 3 became oversensitive, the yellow bands in the test pattern would be displayed as red. On the other hand, if the detector lost sensitivity, the red areas would be displayed as yellow. The changes displayed for fault(s) in the detectors for severity-levels 1 and 2 would be similar.

Transmitter and receiver integrity are monitored by the noise band in the test pattern. If the noise band is not continuous, i.e., it appears broken-up or appears periodically, the local oscillator or transmitter is faulty. If the band is missing completely, the most likely cause is poor receiver sensitivity.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 22-59. Flight Management Computer (Typical)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Other circuits continuously monitor performance, loading, and temperature of system power supplies. A fault resulting in operation outside of preset limits causes a total shutdown because all system power is removed. The only indication is a blank screen. For an overtemperature fault, the system recycles on after several minutes, but for other faults, the system is “latched’’ off and must be manually recycled. Should this occur: 1. Verify that AC and DC circuit breakers are in. 2. Rotate BRT/OFF control to off and, after a few seconds, rotate it to on. Then depress the WX pushbutton. If the fault is transitory in nature, the system will operate satisfactorily; if not, it will automatically cycle off.

FLIGHT MANAGEMENT SYSTEM General The Dash 8 uses a King KNS 660 flight management system (FMS). It is a comprehen sive computer which integrates the use of multiple navigation systems and sensors. Pilot workload is minimized by central programming of the navigation systems and display of information through a central CRT. The system computes great-circle routes, significantly reducing flight time and fuel used over long distances. Figure 22-59 shows one model of the KNS 660 flight management computer. The system can be configured with a variety of sensors. In addition to using VOR/DME information from the NAV receivers, the KNS 660 can use VLF/Omega to provide greater navigational accuracy in areas not covered by VOR/DME navigational aids. Options which can be added to the system include:

In the AUTO/LEG mode of operation, the system is capable of automatically tuning and r e a d i n g s e n s o r s l i k e VO R / D M E a n d VLF/Omega without any input from the pilot. The flight management computer takes the navigation sensor inputs and blends them into a single composite airplane position. Accuracy of this composite position is enhanced by using the best characteristics of each type of sensor. For example, an AHRS has excellent shortterm characteristics while VLF/Omega maintains excellent long-term stability. Additional capabilities of the system include navigation from present position direct to any waypoint, trip plan and fuel plan functions, and the capability of creating a pseudo-VORTAC at any waypoint and establishing an offset parallel course. When one FMS is installed, F M S i n f o r m a t i o n i s s e n t t o t h e p i l o t ’s instruments. If dual systems are installed, FMS information goes to both pilots. For detailed system description and operation, refer to the vendor manual, King KNS 660 Pilot’s Guide. A checklist-form KNS 660 Nav Management System Simplified Operating Instructions is also available. It is concise and broken into the following sections: • Initialization (Turn On) • Page Display Definitions • Direct To Operation • Creating a New Flight Plan • Clearing a Flight Plan/Creating a Ref Waypoint

• Inertial • TACAN • Global positioning system (GPS)

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

KNS 660 Controls The KNS 660 controls are either two-position rocker switches or alternate action push-button switches. Refer to Figure 22-59 for the control positions. The following controls are located on the FMC: • ON–OFF switch—A rocker switch that, when pressed at the top, turns the system on and initiates the self-test. When pushed at the bottom and held for approximately two seconds, the unit turns off. Prior to turning off, a caution message is presented on the screen. • BRT–DIM switch—A rocker switch that, when pressed at the top, increases brightness and message light intensity and decreases it when pressed at the bottom • FRQ key—Selects the two frequency pages which allow frequency management of compatible King avionics • OBS/LEG key—Selects the KNS 660 method of operation. Each key push selects the next method of operation in the sequence of OBS, AUTO/LEG, AUTO 3D, and back to OBS. • SNS key—Selects the active sensor to be used for navigation. Alternate action selects VOR, OMEGA, TACAN, INS, or BLEND, depending upon the sensor configuration of the particular system. • MOD key —Allows selection of NAV, RNV ENR, or RNV APR

• WPT key—This key has two functions: to cycle through the waypoint associated with the active flight plan (FPL 0) and to display the waypoint pages of other waypoints in the system. • NAV key—Using alternate action, selects either NAV 1 or NAV 2 pages • DAT key—Allows viewing of data menu pages (DATA 1 and DATA 2) • HLD key—Allows viewing of the hold pages (HOLD 1 and HOLD 2) • ↑↓ keys (cursor keys)—Used to position the cursor • CLR key—Used to clear a single character in a data field, a complete data field, or an entire page • ENTER key—Used to insert data displayed under the cursor or a complete page of information into the computer memory; also used to select various menu items and to approve specific cursor statements • Alpha-numeric entry keys (KCU 568 CDU)—The KCU 568 has 36 alpha-numeric keys, ten of which are used to enter numerals 0 through 9 and 26 keys which are dedicated to entering the characters A through Z. Eight of the ten numeric keys also can be used to enter north, south, east, west, left, right, minus, and plus.

• MSG key—When pressed, selects the message page, and when pressed again, deselects it • D key (direct to)—Allows selection of DIRECT TO operation • FPL key—Selects active flight plan page (FPL 0) or the flight plan menu (FPLS)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AFCS INTERFACE UNIT (NON-EFIS)

B, C and D. The descriptive titles of these PCBs are as follows: • Dimming and test, PCB—A

DESCRIPTION

• Overtemperature , PCB—B

General

• NAV switching, PCB—C (No. 2 systems)

The AFCS (Automatic Flight Control System) interface unit primarily provides the following functions: • Digital flight data recorder (DFDR), cockpit voice recorder (CVR), and block time recorder (SOO 8027) operational power control • GPWS FAIL annunciator inhibit during stall condition • CVR erase interlock • Elevator trim switching • DME HOLD switching • Switching of radio altimeter information to the ground proximity warning system (GPWS) • Advisory lights bright/dim control • Dimming and test of avionics advisory lights system • Overtemperature indication of AHRS computer (AHRU) • NAV radio switching

• NAV switching, PCB—D (No. 1 systems) The two NAV switching PCBs provide navigation switching function, NAV logic, power and monitoring functions for all radio NAV systems associated with the flight guidance and indicators. The five relays are mounted on the underside of the chassis, and are identified K1 through K5. The two voltage regulators, identified U1 and U2, are mounted on the back of the chassis. The front panel of the AFCS interface unit consists of a toggle switch identified FGC TEST; two pushbutton switches, one identified LT TEST, and the other RESET; five LED indicators in a line below the LT TEST switch (identified SYSTEM 1 at the top for No. 1 systems indicators), and another five LED indicators below the RESET switch identified SYSTEM 2 at the top for No. 2 systems. On the back of the AFCS interface unit, and stacked vertically, are three electrical connectors, identified (from top to bottom) A, B and C. Between connectors B and C are three keyways to ensure correct unit installation. At the base of the unit, at each end, is a guide hole to aid in installation.

• Flight guidance computer test The AFCS interface unit is enclosed in a 3/8 ATR short case and is located in the avionics rack. Mounted on the chassis of the AFCS interface unit are four printed circuit boards (PCBs) containing logic circuits, two PCBs with light emitting diode (LED) indicators, five relays, and two voltage regulators. The four PCBs are identified on the chassis by letter A,

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DFDR, CVR and Block Time Recorder Operational Power Control Operational power control function to the DFDR and block time recorder is provided by DFDR power relay K3, while for the CVR it is provided by relay K4. For details on operational power control to the DFDR and block time recorder refer to Digital Flight Data Recorder System and Digital Clock, Chapter 31, and for the CVR refer to Cockpit Voice Recorder System, Chapter 23.

GPWS Fail Annunciation Inhibit During aircraft stall conditions, the GPWS would normally provide a fail annunciation. To prevent this occurrence, the fail annunciation signal, which is connected through the normally closed “A” contacts of relay K5, is interrupted by K5 being energized from the voltage which powers the stall warning horn.

CVR Erase Interlock Function CVR erase interlock function is provided by relay K7 located on the overtemperature PCB—B (for details refer to Cockpit Voice Recorder System, Chapter 23).

Elevator Trim Switching Function Relays K1 and K2 provide elevator trim switching function by isolating the elevator trim servo from the flight guidance system and connecting the standby elevator trim control system (refer to Chapter 27 for further details).

DME HOLD Switching The dual inline packing (DIP) switches S1 and S2 in the overtemperature PCB—B provide the means of presetting the interface scope to enable DME HOLD at either No. 1 or No. 2 NAV Mode display panel, as required.

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When the aircraft is fitted with two DME systems, the DIP switch settings are reversed to enable DME HOLD at No. 1 NAV Mode display panel for No. 1 DME system and at No. 2 panel for No. 2 DME system. The switches are preset at the factory during installation.

Switching of Radio Altimeter Information to GPWS The DIP switch S3, relay K6, diodes D16 and D17 in overtemperature PCB—B provide switching of radio altimeter information to the GPWS. During normal operation, decision height information from the No. 1 radio altimeter is supplied directly to the GPWS. In a dual radio altimeter installation, this circuit automatically connects the decision height and altitude signals from No. 2 radio altimeter to the GPWS in event of failure in No. 1 radio altimeter system (refer to Ground Proximity Warning System for details). The DIP switch S3 is preset at the factory according to the radio altimeter system installed.

Advisory Lights Bright/Dim Control Relays K16 and K17, and the associated resistor network consisting of resistors R19 and R21 through R23 in each NAV switching PCB, make up the bright/dim control circuits for the advisory lights for the following: • Advisory display panels • AFCS controller • Altitude alerter • Altitude and heading reference system (AHRS) controllers The bright/dim control circuits in NAV switching PCB—D controls the annunciator power to AFCS controller (No. 1 system annunciator), No. 1 AHRS controller, and No. 1 advisory display panel.

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The bright/dim control circuits in NAV switching PCB—C controls annunciator power to the AFCS controller (No. 2 system annunciators), altitude alerter, No. 2 AHRS controller, and No. 2 advisory display panel.

Dimming and Test of Avionics Advisory Lights System Dimming and test PCB—(A) in the AFCS interface unit provide advisory lights bright/dim and test function to the avionics advisory lights system. The system advisory lights that are controlled through this PCB are as follows: • Passenger address and entertainment system (Chapter 23) • Audio integrating system (Chapter 23) • Ground proximity warning system • Radio navigation systems • Flight guidance system • Inertial reference system (if installed) The test function and dim operation is controlled by the TEST ADVSY/CAUT and the DIM/BRT switches, located on the overhead panel, that operate all advisory lights in the flight compartment. The PCB—(A) operates in like manner as the advisory lights DIM/TEST circuit cards in the No. 1 advisory lights control box (Chapter 33).

Overtemperature Indication The AFCS interface unit provide overtemperature indication of the attitude and heading reference units No. 1 and No. 2 (AHRU). Spare circuits are provided for overtemperature indication of other system components. In overtemperature PCB—(B), magnetic latching-type relays K1 and K2, resistors R1 and

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R2, diodes D1 through D4, D11 and D12 make up the overtemperature switching circuits for the components of No. 1 system. Overtemperature switching circuits for the components of No. 2 system are identical, consisting of relays K8 and K9, resistors R6 and R7, diodes D18 through D21, D28 and D29. The LED indicators are on two separate PCBs, one PCB for No. 1 system components and the other PCB for No. 2 system components. All LEDs are mounted to provide indications on the front panel of the interface unit. An overtemperature condition in any of the systems energizes the associated relay, which in turn puts on the LED indicator connected to it. Switch SW2, identified LT TEST, provides a means of testing all overtemperature indicators by applying a ground to energize the magnetic latching relays to put on all overtemperature LED indicators. Switch SW3, identified as RESET, provides the means of resetting all overtemperature indicator circuits by applying a ground to energize the reset coils of the magnetic latching relays.

Nav Switching Circuits The AFCS interface unit provides complete control, signal and monitoring circuits between both No. 1 and No. 2 nav systems, and their respective flight guidance computers. The two switching circuits are identical, contained on PCB—(D) (No. 1 system), and PCB—(C) (No. 2 system). Inputs to the switching circuits are provided from the VHF nav (VOR), auxiliary navigation source, area navigation (RNAV), or microwave landing system (MLS). The last three systems, however, are customer options and are not installed in every aircraft. Signals supplied to the unit and switched, are: • DME distance, clock/sync and valid • Course select X-Y-Z

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• To-from (+ and –) • Bearing X-Y-Z • Heading output and valid • Tuned-to-localizer (TTL) Signal switching is performed by relays, controlled by solid-state buffers and drivers. The switching circuits receive their control information from the flight guidance controller, depending on which nav source is selected. The VOR inputs are wired through normallyclosed contacts in the deenergized position, so that in event of a relay or switching circuit failure, the VHF nav connections to the flight guidance computer are not interrupted. For EFIS-equipped aircraft, the function of NAV switching is accomplished in the symbol generators, not in the AFCS interphone box.

Flight Guidance Computer Test The FGC TEST switch is a double-pole, singlethrow type switch. One pole of the switch is connected to No. 1 FGC, while the other pole is connected to No. 2 FGC (refer to Flight Guidance System for details on test).

TRAFFIC ALERT AND COLLISION AVOIDANCE (TCAS) SYSTEM DESCRIPTION The traffic alert and collision avoidance system is an independent airborne system. It is designed to act as a backup to the air traffic control (ATC) system and the “see and avoid” concept. TCAS consists of four aircraftmounted antennas; a TCAS computer unit and Mode S transponder located remotely; and displays and controls in the cockpit.

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TCAS has a surveillance volume defined by a minimum horizontal radius of 14 nautical miles and a minimum vertical range of ± 2,700 feet. TCAS continually surveys the airspace around an aircraft, seeking replies from other aircraft in the vicinity via their ATC transponders. The transponder replies are tracked by the TCAS system. Flightpaths are predicted based upon these tracks. Flightpaths predicted to penetrate a collision area surrounding the TCAS aircraft are annunciated by TCAS. TCAS generates two types of annunciations: a traffic advisory (TA) and a resolution advisory (RA). The airspace around the TCAS aircraft where a TA is annunciated can be thought of as a caution area. The physical dimensions of the caution and warning areas are time-based and vary as a function of horizontal and vertical closure speeds (range rate and altitude rate) and horizontal and vertical distances (range and altitude) between the TCAS aircraft and the intruder aircraft. The time-based dimensions are calculated by the TCAS computer using the functions range divided by range rate, and altitude divided by altitude rate. A similar time calculation in the vertical direction also must be satisfied in order to generate a TA or RA annunciation. TCAS monitors a time-based dimension of a caution area that extends 35–45 seconds from the time the intruder aircraft is predicted to enter the TCAS aircraft’s collision area. Should an intruder enter the caution area, traffic information in the form of a TA is issued by the system. The traffic displayed includes the range, bearing, and altitude (if available) of the intruder relative to the TCAS aircraft. The flight crew is to use this information as an aid to visually located the intruder in order to avoid a conflict. TCAS also monitors a time-based dimension of a warning area that extends 20–25 seconds from the time at which an intruder would enter the TCAS aircraft’s collision area. Should an intruder enter the warning area, an escape strategy in the form of a RA is issued by the system.

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The RA is a vertical maneuver recommended to the pilot by TCAS in order to increase or maintain vertical separation relative to an intruding aircraft. The RA will be annunciated both visually and aurally. It will consist of either a corrective advisory, calling for a change in aircraft vertical speed, or a preventive advisory, restricting vertical speed changes.

NOTES

TCAS continuously calculates tracked-aircraft projected positions. TAs and RAs are, therefore, constantly updated and provide real time advisory and position information. Once the flightpath of the intruder no longer conflicts with the collision area of the TCAS aircraft, TCAS will announce CLEAR OF CONFLICT to confirm the encounter has ended. The flight crew should then return to the original clearance profile. TCAS generates resolution advisories and traffic advisories against intruder aircraft with ATC transponders replying in Mode C and Mode S. These include altitude in their transmissions. TCAS uses the altitude information for resolution advisory computations. TCAS can generate only traffic advisories against intruder aircraft whose transponders reply in Mode A (nonaltitude reporting).

WARNING TCAS cannot provide an alert for traffic conflicts with aircraft without operating transponders.

TCAS will assist the pilot who, with the aid of the ATC system, has the primary responsibility for avoiding midair collisions.

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D

F

AUTO

ON

G

XPDR

FAIL

IDENT TRAFFIC A T C ON

A

B

1

TCAS TA RA/TA

SBY ATC

1

2

OFF

TST

T C A S

ALT RP

E

C

Figure 22-60. Gables Mode C, Mode S Transponder, and TCAS Control Panel

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Figure 22-60 shows a typical TCAS controller, with a brief description of each function listed below.

A

TCAS/XPDR Mode Selector

Selects the transponder and/or TCAS mode. • TA/RA Mode—The TA/RA mode is the normal operating mode. In the TA/RA mode, TCAS provides traffic advisories and generates resolution advisories. • TA Mode—TCAS provides traffic advisories only; no resolution advisories. This mode prevents TCAS from issuing RAs when the TCAS aircraft is intentionally flying close to another aircraft, i.e., closely spaced parallel approaches. • XPDR ON—Activates the transponder function only; TCAS OFF is annunciated on the TCAS display. • STBY—Select the transponder to standby. TCAS OFF is annunciated on the TCAS display.

B

TEST

Pressing this switch activates the TCAS selftest feature.

C

F

TA DSPLY

• AUTO—Normal TCAS operation switch position that displays RAs or TAs along with proximate traffic when an intruder is detected. • ON—Displays all traffic within range of TCAS or the display, whichever is most restrictive, and ±2,700 feet vertically of own aircraft. • OFF—Displays resolution advisories only.

G

Amber XPDR FAIL Light

Illuminated when selected XPDR has failed. Loss of valid altitude data will also cause TCAS FAIL. It also illuminates with ALT RPTG on and its altitude information source failed. The light can be turned off by selecting an alternate altitude information source or by turning ALT RPTG off.

H

XPDR 1 or 2

Select XPDR 1 or 2 respectively.

NOTES

XPDR Code Window

Displays four-digit XPDR code.

D

IDENT

Same function as conventional XPDR IDENT.

E

ALT RPTG

Same function as conventional XPDR altitude reporting. TCAS is inoperative with ALT RPTG set to OFF.

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1

2

1

4

... .5

2

4

... .5 -10

0

0

6 .5

.5

4

1

-04

1

2

VSI/TRA

6

4 2

VSI/TRA WITH TRAFFIC ADVISORIES TRAFFIC SYMBOLOGY AND DATA TAG

VERTICAL SPEED POINTER (WHITE)

1

2

- SOLID AMBER CIRCLE

RESOLUTION ADVISORY (RA)

- SOLID RED SQUARE

PROXIMATE TRAFFIC

- SOLID CYAN DIAMOND

OTHER TRAFFIC

- HOLLOW CYAN DIAMOND

4

... .5 -05

0 VERTICAL SPEED SCALE (WHITE)

TRAFFIC ADVISORY (TA)

6 .5

1

4 2

2 NM TCAS RANGE RING (WHITE)

AIRPLANE SYMBOL REPRESENTING TCAS EQUIPPED AIRCRAFT (WHITE)

TYPICAL VSI/TRA DISPLAY

Figure 22-61. TYPICAL VSI/TRA Display

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TRAFFIC ADVISORY DISPLAYS

NOTES

The way TCAS traffic is displayed is dependent upon the type of installation. A flat-panel vertical speed/TA/RA (VSI/TRA) display may be used, or TCAS traffic may be presented on EFIS-type displays. Symbology and colors are the same no matter what type of display is used. Differences in operation do exist and those differences are as follows.

FLAT-PANEL VSI/TRA DISPLAY Many aircraft use the combined vertical speed/traffic indicator. In this indicator, the vertical speed indicator (VSI) takes on the additional function of displaying traffic and resolution advisories, in addition to other traffic information designed to improve situational awareness. Internal switching of TCAS automatically presents a TCAS traffic display on the VSI when a traffic advisory is necessary, as long as the TCAS/XPDR mode selector is positioned to TCAS mode, TA or TA/RA, and the traffic switch is in the AUTO position. Full-time display of traffic is available with the traffic switch in the ON position. A white airplane symbol is displayed in the lower center of the VSI representing your TCAS-equipped aircraft. A white range ring made up of 12 dots, each corresponding to a normal clock position, is included. The range ring surrounds the airplane with a radius of 2 nautical miles, and is intended to assist in interpreting TCAS traffic information. The scale of the VSI display is 6.5 nautical miles to the top display edge of the VSI (ahead of the aircraft), 4 miles to the left and right edges, and 2.5 nautical miles to the bottom (rear of your aircraft). Color-coded symbology is used on this display to identify traffic aircraft in your area. Figure 22-61 shows 3 typical TCAS displays. This instrument replaces the conventional IVSI.

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ELECTRICAL POWER SOURCES Avionics power is supplied by 28 VDC and fixed-frequency AC. The DC power is supplied by two batteries, two DC generators, and two TRUs. Three static inverters provide 115VAC, 400-Hz power; they are run by the 28VDC system. Two stepdown transformers are powered from the 115-VAC, 400-Hz buses and provide 26-VAC, 400-Hz power.

MAINTENANCE CONSIDERATIONS GENERAL The use and application of repair materials and general hardware used for maintenance of the airplane are described in Chapter 20, “Standard Practices,’’ of the Maintenance Manual. Included are procedures, practices, and processes that are not specifically covered in other chapters. Information includes tables, charts, illustrations, and technical data to aid in general maintenance of the airplane.

PITOT-STATIC SYSTEM The static system must be absolutely secure since any leakage of cabin pressure into the system will seriously affect the accuracy of the flight instruments. Take care when performing maintenance on or around the pitot-static lines and instruments. Pitot and static lines should be installed for position drainage of moisture from the lines. Position the lines to eliminate traps where moisture can collect. Observe the following cautions when working on the pitotstatic system.

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CAUTION When a pitot or static pressure line is disconnected, a leak check must be performed (check local authority for this).

CAUTION Do not apply pitot or static pressure to the system unless electrical power is on or damage may result to the air data computer.

CAUTION Do not leave pitot-static heaters on for more than one minute on the ground or they may burn out.

AIR DATA SYSTEM The components of the air data system are remove-and-replace items only. Inspections include examining air data components and lines for security, evidence of leakage, cracks, and evidence of overheating. Take care not to bend any tubing runs in such a manner that will cause kinks or flatten the tubing.

AUTOMATIC FLIGHT CONTROL SYSTEM Most of the autopilot/flight director components are remove-and-replace items only. Inspections include examining AFCS components for security, evidence of leakage, cracks, and evidence of chafing. Associated electrical wiring must be inspected for security, chafing, and evidence of heat damage. When installing a servoactuator, cable tension must be measured and adjusted.

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FLIGHT INSTRUMENTS AND NAVIGATION SYSTEM COMPONENTS Most flight instruments and navigation system components are remove-and-replace items only. Care should be taken to follow the procedures listed in the Maintenance Manual for the particular instrument or component when replacing, testing, or calibrating it and any other procedure listed which concerns the item, such as required leak checks listed in the pitot-static section of the book.

COMMUNICATIONS Most of the communications equipment is remove-and-replace items only. In addition to the standard checks for placement and security of wire bundles, static dischargers should be checked for lightning damage and for erosion of the metal tip.

FUNCTIONAL CHECKS Functional checks should be performed by authorized personnel only. Some avionics equipment can and must be functionally checked from the airplane, even though the unit which has been installed was bench-tested in the shop (e.g., COM units). Other items may require that the unit be tested at an authorized repair facility, but not necessarily tested after installation (e.g., air data instruments checked in flight).

malfunction. If a functional test is required, observe all maintenance, operational, and safety precautions. Refer to the avionics wiring diagrams furnished with the airplane to isolate a malfunction.

LIMITATIONS Maintenance limitations are defined by the parameters and tolerances listed in the functional checks of individual avionics components and in the use of approved hardware (sealants, adhesives, etc.) and approved maintenance practices (torquing, repairing fluid lines, etc.) listed in Chapter 20,“Standard Practices,’’ of the Maintenance Manual and the particular system chapter under Maintenance Practices. When functionally checking any airplane system or component, ensure that AFM operational limitations are also observed for the equipment being checked and other equipment/systems operated to support the testing. Maintenance personnel must be aware of AFM limitations which ground the airplane.

NOTES

FAULT ANALYSIS Troubleshooting any airplane system requires a complete understanding of the function and operation of the system. Where like components are installed on the airplane (e.g., left and right systems), troubleshooting is usually typical for both sides and is listed only once in the Maintenance Manual. It may be necessary to perform a functional test to aid in isolating a

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CHAPTER 24 ELECTRICAL POWER SYSTEMS CONTENTS Page INTRODUCTION................................................................................................................. 24-1 GENERAL ............................................................................................................................ 24-3 Main and Auxiliary Batteries ........................................................................................ 24-7 Battery Temperature Monitor System ......................................................................... 24-13 DC POWER SYSTEM ....................................................................................................... 24-17 General ........................................................................................................................ 24-17 Starter-Generators........................................................................................................ 24-17 Generator Control Units (GCUs) ................................................................................ 24-19 DC Contactor Box ....................................................................................................... 24-19 Generator Control Switches ........................................................................................ 24-21 Transformer-Rectifier Units (TRUs) ........................................................................... 24-23 Automatic Bus-Tie Operation and BUS FAULT Protection ....................................... 24-25 BUS Bar Protection Unit (BBPU)............................................................................... 24-27 DC External Power System ......................................................................................... 24-29 DC Power Monitor System ......................................................................................... 24-31 Engine Starting ............................................................................................................ 24-33 AC POWER SYSTEM........................................................................................................ 24-37 Variable-Frequency AC Power .................................................................................... 24-37 AC Generator Control Unit (GCU) ............................................................................. 24-41 Fixed-Frequency AC Power ........................................................................................ 24-45 AC Power Monitor ...................................................................................................... 24-49

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ELECTRICAL SYSTEM CIRCUIT PROTECTION ......................................................... 24-50 MAINTENANCE CONSIDERATIONS ............................................................................ 24-53 Removal/Installation.................................................................................................... 24-53 Battery ......................................................................................................................... 24-53

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ILLUSTRATIONS Figure

Title

Page

24-1

Electrical Power Component Location .................................................................. 24-2

24-2

Electrical System Block Diagram.......................................................................... 24-4

24-3

Battery Component Locations ............................................................................... 24-6

24-4

Battery Venting System.......................................................................................... 24-8

24-5

Battery Temperature Monitor Schematic............................................................. 24-10

24-6

BATTERY TEMPERATURE Panel .................................................................... 24-12

24-7

Battery Power Distribution .................................................................................. 24-14

24-8

Battery Control Switches..................................................................................... 24-14

24-9

DC Power Schematic ........................................................................................... 24-16

24-10

Starter-Generator ................................................................................................. 24-16

24-11

Generator Control Units ...................................................................................... 24-18

24-12

DC Contactor Box ............................................................................................... 24-18

24-13

Typical DC Contactor Box .................................................................................. 24-18

24-14

DC CONTROL Panel .......................................................................................... 24-20

24-15

TRU ..................................................................................................................... 24-22

24-16

TRU Electrical System ........................................................................................ 24-22

24-17

DC Power Schematic ........................................................................................... 24-24

24-18

Bus Bar Protection Unit....................................................................................... 24-26

24-19

Bus Fault Reset Switch and Caution Light.......................................................... 24-26

24-20

External DC Power .............................................................................................. 24-28

24-21

DC Power Monitor Panel..................................................................................... 24-30

24-22

Engine Start Control—Electrical Schematic ....................................................... 24-32

24-23

AC Generator ....................................................................................................... 24-36

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24-24

Variable-Frequency AC Power Distribution ........................................................ 24-36

24-25

AC Contactor Box and Components ................................................................... 24-38

24-26

AC Variable Frequency Generation—Simplified ................................................ 24-40

24-27

AC CONTROL Panel .......................................................................................... 24-42

24-28

AC 400-Hz Power System Component Locations............................................... 24-44

24-29

Fixed Frequency .................................................................................................. 24-44

24-30

Fixed-Frequency AC Power Schematic ............................................................... 24-46

24-31

Fixed-Frequency AC Power Distribution ............................................................ 24-46

24-32

AC Power Monitor............................................................................................... 24-48

24-33

Right DC Circuit-Breaker Panel .......................................................................... 24-50

24-34

Left DC Circuit-Breaker Panel ............................................................................ 24-50

24-35

Variable Frequency and Avionics Circuit-Breaker Panels................................... 24-51

24-36

Typical DC Power Distribution ........................................................................... 24-52

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CHAPTER 24 ELECTRICAL POWER SYSTEMS

G EN PL #1 IL O DC #1 EN G O RV M SE TE T S 1 # Y HO S T T BA

T BA

FF

O

ACEN G

INTRODUCTION The electrical system on the Dash 8 includes both DC and AC systems. The DC system consists of storage, generation, distribution, and system monitoring. The AC system consists of generation, distribution, and system monitoring. Provision is also made for connection of external power while on the ground to power either the DC or the AC system.

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24-1

24-2 DC SYSTEM BUS VOLTS

AC SYSTEM LOAD

MAIN BATT

VOLTS

LOAD R ESS R MAIN

L ESS L MAIN L SEC

GEN L TRU

AUX BATT

GEN R TRU

LOAD

VOLTS

LOAD

VARIABLE FREQUENCY A A RIGHT

INVERTERS AUX SEC

LEFT

PRIM

B

R SEC

B

C TEST

C

TEST

AC CONTROL OFF

OFF AUX BATT

OFF MAIN BATT

OFF

GEN 1

BATTERY MASTER

OFF OFF GEN 1

OFF GEN 2

GEN 2

INVERTERS

OFF MAIN BUS TIE

OFF BUS FAULT RESET

EXTERNAL POWER

EXT POWER

L PRIMARY

R

OFF

SECONDARY OFF AUXILIARY

DC STARTERGENERATOR (LEFT ENG SAME)

AC GENERATOR (LEFT ENG SAME)

TRANSFORMER RECTIFIER UNITS

400 HZ INVERTERS PRIMARY SECONDARY AUXILIARY

CIRCUIT-BREAKER CONSOLES (L AND R)

CIRCUIT-BREAKER CONSOLES BATTERIES

DC CONTACTOR BOX (ON FWD PRESS BLKHD)

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Figure 24-1. Electrical Power Component Location

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GENERAL

NOTES

The electrical system is primarily a 28-VDC system supplying the majority of airplane electrical services. DC power is produced by two engine-driven starter-generators and two transformer-rectifier units (TRUs). Two 24volt nicad batteries support the DC system. A DC external power receptacle is provided in the left side of the nose for the operation of the DC system from a ground power source. Two engine-driven AC generators provide 115-volt, three-phase, variable-frequency AC power to operate anti-icing heaters and electric pumps and to power the TRUs mentioned above. An AC external power receptacle, located in the right nacelle, is also provided for the operation of the AC system from a ground power unit. Since AC power also supplements the DC system through the TRUs, all AC and DC airplane services can be operated from the AC generators or AC external power alone. Fixed-frequency AC power, required for avionics and some instrumentation, is provided by three solid-state inverters that are powered by the 28-volt system. Inverter output is 115 VAC at 400 Hz. Some of the 115-volt inverter output is reduced to 26 volts, 400 Hz AC through two step-down transformers for operation of 26-volt equipment. Figure 24-1 shows the location of electrical components.

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LEFT DC STARTERGENERATOR

RIGHT DC STARTERGENERATOR

MAIN BATT

AUX BATT

LEFT AC GENERATOR

RIGHT AC GENERATOR

L BATT DIST BUS 3Ø 115 VAC

L 115-VAC VAR FREQ BUS

AC EXT PWR

28 VDC

AC CROSSTIE

L MAIN FEEDER BUS

DC BUS TIE

R BATT DIST BUS

R MAIN FEEDER BUS

DC EXT PWR LEFT TRU

3Ø 115 VAC

28 VDC

R 115VAC VAR FREQ BUS

DC EXT PWR

DC BUS TIE

DC BUS TIE

28 VDC

AC EXT PWR

RIGHT TRU 28 VDC

DC BUS TIE

L SEC FENDER BUS

L ESS BUS

L MAIN DIST BUS

R SEC FENDER BUS

BATTERY POWER BUS

R ESS BUS

R MAIN DIST BUS

AUXILIARY INVERTER

L SEC DIST BUS

R SEC DIST BUS

PRIMARY INVERTER

SECONDARY INVERTER 115 VAC 400 HZ

115 VAC 400 HZ

L 115-VAC, 400-HZ BUS

R 115-VAC, 400-HZ BUS

L 26-V TRANSFORMER

R 26-V TRANSFORMER 26 VAC, 400 HZ

L 26-VAC BUS

26 VAC, 400 HZ

R 26-VAC BUS

LEGEND BATTERY POWER

TRU DC POWER

FIXED AC POWER

DC GENERATOR POWER

VARIABLE DC POWER

26 VAC POWER

EXTERNAL DC POWER

Figure 24-2. Electrical System Block Diagram

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Circuit-breaker consoles are installed behind the pilot and copilot seats, on the outboard side. The circuit-breaker panels for the DC main, essential, and secondary buses for the left circuits are mounted on the pilot’s side and those for the right circuits are mounted on the copilot’s side.

NOTES

The consoles each contain three shelves which mount the 400-Hz inverters and other powerrelated equipment. The avionics circuit-breaker panel is mounted above and behind the pilot’s circuit-breaker console, on the rear flight compartment bulkhead at FS X182.00. The 115-VAC variable-frequency circuitbreaker panel is mounted on the copilot’s side in a similar position on the rear flight compartment bulkhead. A block diagram of the electrical system is provided in Figure 24-2.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DC SYSTEM BUS VOLTS

LOAD

MAIN BATT

YEL

EXTC AFT BTL

YEL

FWD BTL

FUEL VALVE OPEN CLOSED GRN

LOAD L ESS L MAIN

A ESS R MAIN

L SEC

GEN L TRU

AUX BATT

R SEC

GEN R TRU

MT

FUEL CLOSED

VALVE OPEN

MT

GRN

BOTH A

TEST

B

FAULT A

FAULT B

FAULT A

LOOP SELECTION

DC CONTROL

ENGINE 1 PULL FUEL OFF

OFF

OFF

AUX BATT

MAIN BATT

OFF GEN 1

BATTERY MASTER

OFF GEN 2

TEST DETECTION

OFF

OFF MAIN BUS TIE

BATTERY TEMPERATURE

OFF BUS FAULT RESET

•

EXTERNAL POWER

MAIN

GRN

AUX

°C

20

40

60

• SENSOR FAIL

ICE PROTECTION

80 TEST

TAIL AIR FRAME AIR FRAME AUTO

OFF

X.110.0

X68.44 (REF)

UPPER PRESSURE BULKHEAD

SHELF Z.112.00 (LEFT SIDE)

DC CONTACTOR BOX WEB Y12.50 LEFT WALL OF WHEEL WELL

BATTERY METERING SHUNTS LOWER PRESSURE BULKHEAD

SUMP JAR MAIN BATTERY AUXILIARY BATTERY—15AH OR—40 AH (SOO 8070)

SHELF Z.97.00 (LEFT SIDE)

Figure 24-3. Battery Component Locations

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

MAIN AND AUXILIARY BATTERIES

NOTES

The main and auxiliary 24-volt nicad batteries are located in the nose compartment on shelf Z97.00, parallel to the airplane centerline and forward of the lower pressure bulkhead (Figure 24-3). Both batteries are secured in their mounting trays by tiedown bolts. The main battery has a 40-amp-hour capacity and the auxiliary battery has a 15-amp-hour capacity. As an option (SOO 8075) a 40-amphour auxiliary battery can be installed in place of the standard auxiliary battery. The two battery cases are commonly vented to the atmosphere through a sump jar. The jar is mounted aft of the frame at FS X68.44, left side, and is connected to the battery cases and outside air by hoses and a tee union. The negative side of each battery is connected through separate battery metering shunts to airframe ground on the underside of shelf Z112.00 above the batteries. Each battery is equipped with temperature sensors which are connected to the temperature monitoring system. The battery contactors, control relays, and circuit breakers are located in the DC contactor box, which is mounted on the forward side of the upper pressure bulkhead.

CAUTION If CSI 82066 is installed and the batteries are selected on with external DC power applied, the battery charge rate and temperature must be continually monitored to ensure that battery temperature limits are not exceeded.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 24-4. Battery Venting System

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Battery Venting System

NOTES

The battery venting system (Figure 24-4) consists of the battery sump jar, breather tubes from the two batteries, and the overflow vent through the airplane skin to the atmosphere. The sump jar is made of styrene plastic and consists of two separate parts. The top is mounted to the left wheel well web forward of the main battery and contains two vent pipe fittings, one connected to the battery breather tube and the other to the overflow vent. The bottom part of the sump jar screws into the top and is held in place by two halfclamps, also attached to the wheel well web. The jar contains a pad soaked with a boric acid solution which neutralizes any fumes or electrolyte spillovers from the batteries. The batteries are each equipped with two vents, but one is capped and unused. The breather tubes from the batteries are plastic tubes joined at a tee connector, which in turn is connected to a fitting on top of the sump jar. The tubes are clamped at various places to prevent movement and/or damage. The second vent fitting on the sump jar is connected through a plastic tube, the end forming the overflow through the airplane skin on the bottom of the nose section. It is clamped to prevent movement, and the end is cut flush with the airplane skin.

NOTE It is recommended that the battery venting system be checked at every battery removal and the sump jar cleaned as required.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

BATTERY TEMPERATURE MAIN 20

°C

40

60

80

AUX

TEST

AUX BAT R HOT

R

SENSOR FAIL A B C

P

J

H G

F

R

R

MAIN BAT HOT

LOGIC E

L

M

P/O WARNING PANEL

N R 10

11

BATT TEMP IND (L8) 6 CB2 BATT TEMP CAUT LTS (M8)

28-VDC R ESS BUS

5 CB1

HI

5-VDC OVERHEAD PANEL LIGHTING

LO

F

C

B

E

E

T

B

C

F

T

T

DISPLAY SENSOR

T

OVERHEAT SENSOR LEFT (AUX) BATTERY

RIGHT (MAIN) BATTERY

Figure 24-5. Battery Temperature Monitor Schematic

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Temperature Sensors

NOTES

There are two independent temperature sensors (thermistors) in each battery. Each thermistor is mounted on an intercell connector link of the battery, with external connections provided through a six-pin connector on the battery case. One sensor (display sensor) provides the temperature input to the related indicator driver circuit, and the other (overheat temperature sensor) to the overheat warning circuit in the monitor. The sensor thermistors vary their internal resistance nonlinearly with changes in temperature ranging from 31,439 ohms at 24° C (75.2° F) to 3,251 ohms at 80° C (176° F), with a tolerance of ±5%. A threeposition test switch is schematically shown in Figure 24-5.

NOTE If the airplane has recently been operating or sitting in the hot sun, the battery temperature indication may be higher. If ambient temperature is below 15° C (59° F), the first green segment remains on, indicating a serviceable system.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

MAN BAT HOT

AUX BAT HOT

WARNING LIGHTS ILLUMINATE TO INDICATE THAT TEMPERATURE OF RELATED BATTERY HAS EXCEEDED 65° C.

MAIN BATTERY OVERHEAT SENSOR FAIL LIGHT ILLUMINATES TO INDICATE FAILURE OF MAIN BATTERY OVERHEAT SENSOR

BATTERY MONITOR TEST SWITCH THREE-POSITION SWITCH (PRESS, PUSH LEFT, AND PUSH RIGHT) PROVIDES THREE TEST SIMULATIONS: • PRESSING IN SIMULATES 70° C TEMPERATURE IN BOTH BATTERIES, CAUSING MAIN AND AUX SCALES TO SHOW 70° AND ILLUMINATING BAT HOT CAUTION LIGHTS. • PUSHING LEFT SIMULATES SHORTED CIRCUITS ON ALL SENSORS, CAUSING SENSOR FAIL LIGHTS TO ILLUMINATE AND SINGLE YELLOW LIGHT SEGMENT AT THE 60° POINT ON BOTH SCALES TO ILLUMINATE. BATTERY TEMPERATURE DISPLAYS SCALES EACH SCALE IS A ROW OF VERTICAL BAR LIGHT SEGMENTS WHICH ILLUMINATE IN LINE FROM LEFT TO RIGHT UP TO THE APPROPRIATE POINT ON THE NUMBERED SCALE TO INDICATE RELATED BATTERY TEMPERATURE. RANGES: NORMAL OPERATION (GREEN LIGHTS)—15 TO 50° C CAUTIONARY (AMBER LIGHTS)—50 TO 65° C OVERHEAT (RED LIGHTS)—65 TO 80° C

• PUSHING RIGHT SIMULATES OPEN CIRCUITS ON ALL SENSORS, GIVING SAME INDICATIONS AS ABOVE. AUXILIARY BATTERY OVERHEAT SENSOR FAIL LIGHT ILLUMINATES TO INDICATE FAILURE OF AUXILIARY BATTERY OVERHEAT SENSOR

NOTE: FAILURE OF A DISPLAY SENSOR IS INDICATED ON THE ASSOCIATED SCALE BY ILLUMINATION OF A SINGLE AMBER SEGMENT AT THE 60° POINT ON THE SCALE.

Figure 24-6. BATTERY TEMPERATURE Panel

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NOTE

BATTERY TEMPERATURE MONITOR SYSTEM The battery temperature monitor system provides continuous temperature indications of the main and auxiliary batteries plus warning of high temperature conditions. A high temperature indication enables action to be t a ke n t o i s o l a t e t h e b a t t e r y a n d p r eve n t p o s s i b l e d a m a g e a r i s i n g f r o m ex c e s s ive battery temperature. The system consists of a temperature level (display) sensor, an overheat sensor (installed on the intercell links of each battery), a BATTERY TEMPERATURE panel (Figure 24-6) on the overhead console, and two warning lights (Appendix B). The MAIN and AUX battery temperature display scales on the BATTERY TEMPERATURE panel indicate battery temperature, as sensed by their associated display sensors. The displays use luminous green, amber, and red light bar segments through a range of 15 to 80° C (59 to 176° F). MAIN BAT HOT and AUX BAT HOT warning lights, located on the warning lights panel, are illuminated by their respective overheat sensors when sensed temperature exceeds 65° C (149° F). A failure warning mode is incorporated into the display sensor circuits, which warn of display sensor failure (in the form of shorted or open sensor circuits) by illuminating a single amber light segment at the 60° C (140° F) point on the associated temperature scale. Failure warning of the main or auxiliary battery overheat sensor is provided by SENSOR FAIL lights on the BATTERY TEMPERATURE panel, which illuminate upon detection of shorted or open overheat sensor circuits.

Mod 8/1487 introduces a new battery temperature monitor vendor, (TM O R S ) . T h i s n ew m o n i t o r h a s different test indications than the old (Weston) type. These are:

1.

Press test button: • MAIN and AUX displays both show 70 °C (±2.5 °C) and both warning lights and master warning light illuminate.

2.

Press the test button to the left and hold: • Both warning lights and master warning light illuminate, both FAIL lights and 60° C elements are lit, and both main and aux rows of elements are lit sequentially from normal (local ambient) temperature to full scale, and then turned off. Release the test button: • Check that both FAIL light, 60° C elements and warning lights are turned off and all elements are lit simultaneously followed by a progressive turn off from full scale to normal temperature.

3.

Press the test button to the right and hold: • Check that both fail lights turn on and the lit elements showing the local ambient temperature turn off progressively until no LEDs are lit and that the 60° C elements are lit. Release the test button.

Incorporated into the monitor system is a threeway test function which permits test simulations of overtemperature conditions, sensor short circuit failure modes, and sensor open circuit failure modes. The test uses a single battery monitor TEST switch on the BATTERY TEMPERATURE panel.

Revision 2

• Check that both FAIL lights are turned off and the elements are lit progressively from off scale at low end to local ambient and the 60° C elements are turned off.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

R MAIN DIST BUS BATTERY L MAIN DIST BUS NOTE: POWER BUS 2 CBs INSTALLED CR3 CR2

FLIGHT COMPARTMENT

R SEC DIST BUS

L SEC DIST BUS L ESS BUS

R ESS BUS

LEFT (NO. 1) GENERATOR

CR5

RIGHT (NO. 2) GENERATOR

CR4

K1

K2

LEFT MAIN FEEDER BUS

K5

K7

K9 F1

AC VAR

RIGHT MAIN FEEDER BUS K6

K22

L SECONDARY FEEDER BUS K17 DC CONTACTOR BOX ASSEMBLY

K21 CR1

K3

F5

K4

K8

R SECONDARY FEEDER BUS K10

K18 F4

F3

+ LEFT TRU

F2 +

L BATT BUS

+

+

LEFT BATTERY

R BATT BUS

RIGHT BATTERY

RIGHT TRU

AC VAR

EXTERNAL POWER RECEPTACLE

+

Figure 24-7. Battery Power Distribution

MAIN BATTERY SWITCH

DC CONTROL

OFF

AUXILIARY BATTERY SWITCH

AUX BATT

AT AUX BATT, LEFT MAIN FEEDER BUS CHARGING CIRCUIT IS CONNECTED TO AUXILIARY BATTERY IF BATTERY MASTER SWITCH IS SELECTED

GEN 1

OFF MAIN BATT

OFF

OFF BATTERY MASTER

OFF GEN 2

AT MAIN BATT, MAIN BATTERY IS CONNECTED TO RIGHT MAIN FEEDER BUS IF BATTERY MASTER SWITCH IS SELECTED.

OFF

OFF

MAIN BUS BUS FAULT EXT TIE RESET POWER

BATTERY MASTER SWITCH AT BATTERY MASTER, BOTH BATTERIES ARE CONNECTED TO ESSENTIAL BUSES.

Figure 24-8. Battery Control Switches

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Battery Control

NOTES

The batteries are connected to the essential buses and main feeder buses by contactor relays (Figure 24-7), which are controlled by switches on the DC CONTROL panel in the flight compartment (Figure 24-8). A single lever-locked BATTERY MASTER switch connects both batteries to the essential buses through K3 and K4, and separate AUX BATT and MAIN BATT switches control connection of the associated battery through K7 and K8, to the related feeder bus, (the BATTERY MASTER switch must be on to connect the batteries to the feeder buses). MAIN BATTERY and AUX BATTERY caution lights illuminate whenever the related battery is not connected to its feeder bus (Appendix B). A lever-locked MAIN BUS TIE switch on the DC CONTROL panel provides for manual tying of power from the main battery through the right main feeder bus, through K21, to the left main feeder bus (Figure 24-8).

NOTE Battery power cannot be applied to the secondary buses.

Revision 4

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

R MAIN DIST BUS BATTERY L MAIN DIST BUS NOTE: POWER BUS 2 CBs INSTALLED CR3 CR2

FLIGHT COMPARTMENT

R SEC DIST BUS

L SEC DIST BUS L ESS BUS LEFT (NO. 1) GENERATOR

R ESS BUS CR5

RIGHT (NO. 2) GENERATOR

CR4

K1

K2

LEFT MAIN FEEDER BUS

K5

K7

K9 F1

AC VAR

RIGHT MAIN FEEDER BUS K6

K22

L SECONDARY FEEDER BUS K17 DC CONTACTOR BOX ASSEMBLY

K21 CR1

K3

F5

K4

K8

R SECONDARY FEEDER BUS K10

K18 F4

F3

+ LEFT TRU

F2 +

L BATT BUS

+

LEFT BATTERY

+

R BATT BUS

RIGHT BATTERY

RIGHT TRU

AC VAR

EXTERNAL POWER RECEPTACLE

+

Figure 24-9. DC Power Schematic

Figure 24-10. Starter-Generator

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

DC POWER SYSTEM

in each brush block by springs. A removable window strap provides access to the brushes.

GENERAL

An air inlet duct is secured to the nondriven end of the starter-generator and an exhaust duct is incorporated in the mounting flange to provide cooling. A thermostatic switch is installed in the starter-generator to close a circuit if temperature exceeds a predetermined limit and provides an appropriate DC GEN HOT indication on the caution lights panel.

DC power output (Figure 24-9) from the starter-generators and TRUs is distributed via feeder buses to a series of distribution buses designated main, essential, and secondary. Generator output is normally applied through main feeder buses to the main buses and from the main buses to the essential buses. TRU output is normally applied through secondary distribution buses. Provision is made to automatically interconnect the main and secondary feeder buses, thereby compensating for the loss of up to two power sources without affecting the supply of DC power to the buses.

Each generator is controlled by its individual GCU so that the output voltage remains constant over the speed range of approximately 5,600 to 12,000 rpm.

NOTES

STARTER-GENERATORS The two starter-generators, one mounted on the accessory gearbox of each engine, serve as starter motors and automatically revert to generator operation following a successful engine start. Generator control units, designated GCU No. 1 (left) and GCU No. 2 (right), regulate starter-generator output to the associated main feeder buses through main bus contactors. The start mode operation of the startergenerators is controlled by the engine start circuits (MSM Chapter 24) in conjunction with the GCUs.

Starter-Generator Operation The generators are rated to supply 200 amperes at 30 volts when self-cooled and 300 amperes at 30 volts when air-blast cooled. The starter-generator is a four-pole, shunt-connected, fully compensated DC design with interpole windings, brush commutation, and internal four-blade aluminum die-cast cooling fan, and four brush blocks, spaced 90° apart around the commutator. A pair of brushes with leads joined to a common terminal are retained

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

STARTER GENERATOR − MAINTENANCE PRACTICES

2. Fully engage starter−generator drive; engage dowel pin with dowel hole.

Removal/Installation

3. Position clamp on starter−generator/adapter joint, ensuring complete contact. Tighten clamp nut and torque 55 to 60 pound/ inches.

A. Remove Starter−Generator 1. Remove power from aircraft buses 2. Open engine cowl center panel 3. Disconnect electrical plug. Blank connector plug and associated receptacle on starter−generator.

CAUTION Ensure starter−generator is suitably supported, parallel to engine horizontalcenterline, during disengagement of starter−generator drive, to prevent possible damage to accessory drive components.

4. Support starter−generator; remove clamp and withdraw starter−generator from mounting flange and transition duct. Remove and discard preformed packing B. Install Starter−Generator 1. Lubricate new preformed packing and starter−generator drive splines with engine oil of correct specification.

CAUTION 1. Proper torque must be applied to prevent damage from arcing or overheating. 2. Overtorque may cause cracking of the terminal block or stripping of the threads. 3. The specified torque values given below are to be strictly adhered to

4. Remove insulation from cable ends. Install cable on starter−generator terminal (larger) positive stud and cable on starter−generator terminal (smaller) negative stud. Secure with new locknuts and washers.

NOTE Correct polarity of starter−generator cables is provided by hole size of cable ends. Positive cable hole is of increased diameter.

5. Reinstall terminal block cover oriented as shown; secure with screws.

CAUTION Ensure starter−generator is suitably supported, parallel to engine horizontal centerline, during engagement of starter−generator drive, to prevent possible damage to accessory drive components.

TR # 24-1—March 2010

TR24.1-1

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

C. With GCU-2 energized, start contactor 2431-K2 energizes to connect the starter section of the No. 2 engine starter-generator to the right main feeder bus and GCU-2 to the 28-VDC right essential bus. The starter-generator, driving through the turbomachinery accessory gearbox, rotates No. 2 engine highpressure compressor. Fuel is switched on at a predetermined starter-generator speed (10%to 19%Nh), and light-up takes place. Startergenerator speed is sensed by the speed sensor (magnetic pickup) in the starter-generator, which relays speed signals to GCU-2. For a detailed schematic analysis of engine start, refer to MSM Chapter 24.

TR24.1-2

NOTES

TR # 24-1—March 2010

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 24-12. DC Contactor Box K19 K22 K5

K17

K9

K1

K7

K3 K26

Figure 24-11. Generator Control Units K20 K2

K6

K18

K21

K23

K10

K4 K8

RELAYS AND FUNCTIONS (* INDICATES ON RELAYS PANEL, NOT IN DC CONTACTOR BOX) K1

#1 GEN CONTACTOR

K14

R TRU LOGIC *

K2

#2 GEN CONTACTOR

K15

EXT PWR INTERLOCK RELAY

K3

AUX BATT—L ESS BUS

K16

EXT PWR TIME DELAY

K4

MAIN BATT—R ESS BUS

K17

L TRU/L SEC BUS

K5

L MAIN/L SEC BUS TIE

K18

R TRU/R SEC BUS

K6

R MAIN/R SEC BUS TIE

K19

L TRU UNDERVOLTAGE

K7

AUX BATT/L MAIN FEEDER

K20

R TRU UNDERVOLTAGE

K8

MAIN BATT/R MAIN FEEDER

K21

MAIN BUS TIE

K9

L EXT PWR

K22

SEC BUS TIE

K10

R EXT PWR

K23

EXT PWR OVERVOLTAGE

K11

L GEN LOGIC *

K24

EXT PWR CHANGEOVER*

K12

R GEN LOGIC *

K25

EXT PWR CHANGEOVER *

K13

L TRU LOGIC *

K26

EXT PWR LOCKOUT

Figure 24-13. Typical DC Contactor Box

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GENERATOR CONTROL UNITS (GCUS) The GCUs (Figure 24-11) are hard-mounted by four mounting bolts in the electrical equipment rack under the right DC circuit-breaker panel beside the copilot’s seat. A formed aluminum case houses the circuitry, which, with the exception of the control relays, is all solid state. Electrical connections are via a multipin connector on the side of the unit. When the engines are running (start circuits released) and the generators are switched on, each GCU connects its generator to the main feeder bus and regulates generator output at 28 VDC, regardless of changes in speed and load. The GCU monitors generator input current and output current (using current transformers), interpole current, output voltage, and generator speed. By varying a generator field strength the GCU maintains a constant generator output voltage.

The left and right contactors are physically located on opposite sides of the contactor box assembly and are connected by bus bars mounted directly on the contactors. Two current transformers are installed in each generator system. One is installed in the positive feeder cable adjacent to the generator contactor in the DC contactor box, and the second transformer is mounted on the negative feeder cable connecting the generator E terminal to system ground. A terminal block is secured to the top of the case to provide terminal connections for the transformer secondary, leading to the GCU. A removable plastic cover plate provides protection over the terminals. For a detail of component layout of the DC contactor box, refer to Figure 24-13.

NOTES

The GCUs also monitor the associated startergenerator for fault conditions and automatically shut a generator down and remove it from the feeder bus upon detection of a malfunction. Time delay circuits prevent nuisance tripping due to normal switching transients or line noise and enable the GCU to verify before taking corrective action. The GCU also provides a paralleling control to equalize the load on both generators when operating with the main feeder bus tie closed. The GCU performs current limiting to prevent excessive current from flowing during starting of the opposite engine and field weakening to reduce starter field current as engine speed increases during engine starting.

DC CONTACTOR BOX The DC contactor box (Figure 24-12), located in the airplane nose, contains all primary contactors, the main and secondary feeder buses, bus-tie contactors, protective fuses, and circuit breakers for the main and secondary dist r i bu t i o n bu s e s a n d va r i o u s c o n t r o l a n d interlocking circuits.

Revision 4

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

BUS FAULT RESET SWITCH WHEN MOMENTARILY POSITIONED TO BUS FAULT RESET, THE BBPU CONTROL OVER BUS ISOLATION FOLLOWING A BUS FAULT IS RELEASED AND THE BBPU IS REARMED

EXTERNAL POWER SWITCH WHEN POSITIONED TO EXT PWR, A DC POWER SOURCE IS CONNECTED TO THE LEFT AND RIGHT MAIN FEEDER BUSES. AT THE SAME TIME THE BATTERIES ARE DISCONNECTED FROM THE BUSES.

GENERATOR CONTROL SWITCHES WHEN POSITIONED TO GEN 1 OR 2, THE GCU IS ARMED TO ACTIVATE ITS STARTERGENERATOR TO THE GENERATING MODE. IN OFF, THE GENERATOR IS SHUT DOWN OR, FOLLOWING SHUTDOWN BY OTHER MANUAL OR AUTOMATIC MEANS, RESETS THE GCU FOR SUBSEQUENT REACTIVATION.

EXTERNAL POWER ADVISORY LIGHT ILLUMINATES WHEN EXTERNAL POWER IS AVAILABLE TO THE DC BUSES.

MAIN BUS TIE SWITCH LEVER-LOCKED TO OFF. WHEN POSITIONED TO MAIN BUS TIE, THE LEFT AND RIGHT MAIN FEEDER BUSES ARE MANUALLY TIED. THE BBPU INHIBITS BUS-TIE OPERATION ON DETECTION OF A BUS FAULT.

Figure 24-14. DC CONTROL Panel

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

GENERATOR CONTROL SWITCHES

NOTES

Generator mode operation is controlled by the GEN 1 and GEN 2 switches located on the DC CONTROL panel (Figure 24-14). The switches provide for normal activation and deactivation of the associated generator through its GCU. The switches also have a GCU reset function, by which a starter-generator that has been automatically tripped off line due to a momentary fault can be reactivated by positioning the related generator control switch to OFF and then back on. The No. 1 DC GEN and No. 2 DC GEN caution lights, located on the caution lights panel (Appendix B), illuminate in response to signals from the GCUs and indicate that the related starter-generator is off line. The No. 1 DC GEN HOT and No. 2 DC GEN HOT caution lights are also provided to warn of gene r a t o r ove r h e a t . E a c h i s i l l u m i n a t e d b y overtemperature sensors in the associated starter-generator when internal temperature exceeds a predetermined limit. The light extinguishes when temperature drops below the overheat point.

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 24-15. TRU

CR3

L ESS

CB32 CB34

K3

L BATT BUS

CB75

K5

CB33

CB21 CR2 R MAIN

K6

CB35

K4

P/O S7 BATT MASTER

K 21

P/O CAUTION PANEL L TRU FAIL LOGIC CIRCUIT

CB13

CR3 L MAIN

CB6 P/O S7 BATT MASTER

CR4

R ESS

R BATT BUS

P/O CAUTION PANEL R TRU FAIL

28V

LOGIC CIRCUIT

28V

CB36

X2

X1

UNDER CR10 VOLTAGE RELAY

CR9

I

12

13 14

F

C1

B3

B3

B2

B1 A3

B1 A3

A2

A1

A1 X1 K14 R TRU LOGIC X2 RELAY

A PH

C1

X1 L TRU X2 LOGIC RELAY

B

C2

CB39

TO A2 CB26 R ESS 28V K13

C3

CB41

B2

C3

B PH

CB42

CB40

115/200V LEFT VAR FREQ BUSES

D C2

CB43

TO 28V BUS TIE BBPU

+ RIGHT TRU

20

D

C PH

B PH

A PH

|

20 CB38

250

K20 X1

C PH

250

UNDER VOLTAGE RELAY

B1

X2

K19

B

A1 A2

B2

F2 225 A

CR7

B3

SHUNT R6

CR8

+ | LEFT TRU I

11

K18 CB37

B3 B2

SHUNT R5

X2

B3

B1

F

X1

K22 B2 P/O K12

F1 225 A

R SEC

150 A

X2

K17

L SEC

X1

A2

12 11

A1

14

13

F5

115/200V RIGHT VAR FREQ BUSES

Figure 24-16. TRU Electrical System

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TRANSFORMER-RECTIFIER UNITS (TRUS)

NOTES

The two TRUs (Figure 24-15), designated left and right, consist of voltage transformers and diode rectifiers. Both identical TRUs are installed in the nose compartment on the left side of a shelf at Z112.00. Each TRU is fed with three-phase, 115-volt, variable-frequency AC power through a circuit breaker from the related variable-frequency AC bus. The AC input is converted to a nominal 28-VDC output.

NOTE Actual TRU output voltage can normally range from 25 to 29.5 volts, depending on secondary DC load with a maximum current loading of 200 amperes.

DC power is normally available from the TRUs whenever there is AC generator or AC external power on the related variable-frequency bus. Manual shutdown of a TRU without removing AC power is possible only by pulling the related L TRU or R TRU circuit breaker on the 115-VAC circuit-breaker panel (Figure 24-36). TRU connection to the secondary feeder buses is controlled automatically by TRU undervoltage relays which operate secondary bus contactors. Each undervoltage relay senses output voltage from its TRU and connects the TRU to the secondary bus whenever this voltage is greater than 18 volts (and subsequently disconnects the TRU if output voltage drops below 18 volts) (Figure 24-16). The L TRU and R TRU caution lights on the caution lights panel illuminate in response to signals from the undervoltage relays to indicate that the associated TRU is off line. Warning of TRU overheat is provided by the illumination of L TRU HOT and R TRU HOT caution lights (Appendix B) in response to overtemperature sensors in each TRU. The TRU HOT light extinguishes when temperature drops below the overheat point.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT

L MAIN DIST BUS

R MAIN DIST BUS BATTERY POWER BUS

NOTE: 2 CBs INSTALLED CR3

CR2

L SEC DIST BUS

R SEC DIST BUS L ESS BUS

R ESS BUS

CR5

LEFT (NO. 1) GENERATOR

CR4 RIGHT (NO. 2) GENERATOR

K1

K2

K21

LEFT MAIN FEEDER BUS

K5

K22

L SECONDARY FEEDER BUS K17

DC CONTACTOR BOX ASSEMBLY AC VAR

RIGHT MAIN FEEDER BUS

CR1

K7

K9

F1 + LEFT TRU

K3

K6

R SECONDARY FEEDER BUS

F5

K4

K8

K10

K18

F4

F3

F2 +

L BATT BUS

+ LEFT BATTERY

+

R BATT BUS

RIGHT BATTERY

RIGHT TRU

AC VAR

EXTERNAL POWER RECEPTACLE +

Figure 24-17. DC Power Schematic

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AUTOMATIC BUS-TIE OPERATION AND BUS FAULT PROTECTION

NOTES

Four bus-tie relays, consisting of a main bus tie, a secondary bus tie, and two main/secondary bus ties, are controlled automatically by a series of DC logic relays that tie the appropriate feeder buses when there is an inoperative power source(s) (the main bus tie is also manually operable). The DC logic relays receive on-line signals from the GCUs and TRU undervoltage relays. These signals are used to determine which bus tie to close in order to sustain power on the affected feeder bus following failure of a power source (Figure 24-17). When only one DC generator is on line, the logic relays close the main bus tie to maintain power to the affected main feeder bus from the opposite generator. The same happens through the secondary bus tie when only one TRU is on line. If both starter-generators are off line and both TRUs are on, the main/secondary bus ties are closed to permit the TRUs to power the main and essential buses. Con-versely, if both generators are on and both TRUs off, the main/secondary bus ties are closed to supply the secondary buses from the starter-generators.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 24-18. BUS Bar Protection Unit

BUS FAULT RESET SWITCH (MOMENTARY SELECTION) DC CONTROL

OFF AUX BATT

MAIN BATT

OFF GEN 1

OFF

BUS FAULT RESET—RELEASES BBPU CONTROL OVER A BUS ISOLATION (FOLLOWING A BUS FAULT) ONLY IF THE FAULT HAS CLEARED. RE-ARMS BBPU IN THE EVENT OF ANOTHER BUS FAULT

OFF

BATTERY MASTER OFF

OFF

GEN 2 MAIN BUS TIE

OFF BUS EXTERNAL FAULT POWER RESET

(OVERHEAD PANEL)

DC BUS

CAUTION LIGHT (AMBER) ILLUMINATED—BBPU HAS DETECTED A BUS FAULT ON MAIN DC BUSES, AND HAS INHIBITED THE BUS TIE CONTACTOR FROM CLOSING

Figure 24-19. BUS Fault Reset Switch and Caution Light

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BUS BAR PROTECTION UNIT (BBPU) Automatic protection for the starter-generators and batteries from overloads due to bus faults (short circuits) is provided by a bus bar protection unit (BBPU) (Figure 24-18) acting in conjunction with the GCUs. The unit is located on the top shelf of the right circuit-breaker console. The BBPU detects bus faults by monitoring the GCUs for generator overcurrent conditions and responds to a fault in two stages. Upon initial detection of a fault, the BBPU immediately opens or inhibits the main bus tie and the two main/secondary bus ties to ensure that the faulty bus is isolated from the rest of the system. At the same time, a DC BUS caution light on the caution lights panel (Appendix B) illuminates to warn of the fault condition.

off-on selection with the appropriate generator control switch (Figure 24-19).

NOTES

If after approximately seven to ten seconds the generator overload is still present, the BBPU trips the affected generator off line and isolates the appropriate battery from the faulted main feeder bus (MAIN or AUX BATTERY and related DC GEN caution lights illuminate). The BBPU continues to monitor the unfaulted side. All main DC services on the faulted side (and all secondary DC services if both TRUs are off line) are lost. Essential bus power from the remaining generator or the batteries is normally unaffected.

NOTE Manual operation of the main bus tie through the MAIN BUS TIE switch is inhibited once the BBPU has reacted to a fault.

If the fault is subsequently cleared, normal operation can be restored by means of a momentary BUS FAULT RESET switch located on the DC CONTROL panel. When the BUS FAULT RESET switch is held, BBPU authority over bus isolation and generator shutdown is canceled (the BBPU remains armed in case the fault returns). Reactivation of the affected starter-generator requires resetting its GCU by

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT COMPARTMENT

L MAIN DIST BUS

R MAIN DIST BUS BATTERY POWER BUS

NOTE: 2 CBs INSTALLED CR3

CR2

L SEC DIST BUS

R SEC DIST BUS L ESS BUS

R ESS BUS

CR5

LEFT (NO. 1) GENERATOR

CR4 RIGHT (NO. 2) GENERATOR

K1

K2

K21

LEFT MAIN FEEDER BUS

K5

K22

L SECONDARY FEEDER BUS K17

DC CONTACTOR BOX ASSEMBLY AC VAR

RIGHT MAIN FEEDER BUS

CR1

K7

K9

F1 + LEFT TRU

K3

K6

R SECONDARY FEEDER BUS

F5

K4

K8

K10

K18

F4

F3

F2 +

L BATT BUS

+ LEFT BATTERY

+

R BATT BUS

RIGHT BATTERY

RIGHT TRU

AC VAR

EXTERNAL POWER RECEPTACLE +

Figure 24-20. External DC Power

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DC EXTERNAL POWER SYSTEM The external DC power source takes precedence over airplane electrical power to supply the DC bus system. When external DC power is applied and selected on, the control circuits provide isolation of the batteries and generators from the external power source. A customer option (CSI 82066) permits the batteries to remain on line for charging from the external DC source. Overvoltage protection from the external source is also provided, as well as an advisory light for visual indication that external DC power is selected.

CAUTION If CSI 82066 is installed, and the batteries are selected on with external power applied, the battery charge rate and temperature must be continually monitored to ensure temperature limits are not exceeded.

When the EXT PWR switch is in OFF, the interlock relays ensure that the batteries are reconnected to the feeder buses before the external power contactors open. In addition, the main bus tie and secondary/main bus ties are held for a further .5 second after external power removal by the changeover relay time delay logic and eliminate any gaps in the power supplied to the main buses during transition to generator power. Protection for the DC power system from external power unit faults is provided by an overvo l t a g e r e l a y. T h e r e l a y a u t o m a t i c a l l y disconnects external power from the system when input voltage at the receptacle exceeds 31.5 volts. A latching relay holds external power out until reset is accomplished by cycling the EXT PWR switch to OFF and then on. This action releases the latching relay and reconnects external power to the buses (Figure 24-19).

NOTES

The DC external power system provides for the connection of a DC ground power unit to the main and secondary feeder buses through a receptacle in the left side of the nose. With a DC power source connected to the nose receptacle and the EXT PWR switch on the DC CONTROL panel switched on, power is applied to the main feeder buses. External power causes the main bus tie and the secondary/main bus ties to close, connecting the secondary buses. An external power interlock relay isolates the batteries from the main feeder buses. A green external power advisory light illuminates when the contactors close to show that external power is available to the DC buses. Since the external control circuits operate from the essential buses, the BATTERY MASTER switch should be on before external power is connected. The generator mode operation of the starter-generators is inhibited when external DC power is applied to the system (start mode operation is unaffected) (Figure 24-20).

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BUS VOLTAGE DISPLAY LUMINOUS DIGITAL INDICATOR DISPLAYS VOLTAGE OF BUS SELECTED ON BUS VOLTAGE DISPLAY SELECTOR SWITCH.

BATTERY LOAD DISPLAYS LUMINOUS DIGITAL INDICATORS FOR EACH BATTERY DISPLAY CHARGE/DISCHARGE RATES AS A DECIMAL FRACTION OF MAXIMUM RATE CAPACITY. EXAMPLES: RATED LOAD IS EXPRESSED AT 1.00; 60% OF RATED LOAD IS EXPRESSED AS .60. DISCHARGING IS INDICATED BY FLASHING MINUS SIGN (—) PRECEDING NUMBER. BUS VOLTAGE DISPLAY SELECTOR SWITCH SELECTS BUS (ESSENTIAL, MAIN, OR SECONDARY) TO BE VOLTAGE-MONITORED ON BUS VOLTAGE DISPLAY. DC MONITOR SYSTEM TEST BUTTON WHEN PRESSED AND HELD, ALL VOLTS AND LOAD DISPLAY SEGMENTS ILLUMINATE FOR TWO SECONDS, FOLLOWED BY APPEARANCE OF THE FOLLOWING TEST VALUES: LOAD DISPLAYS SHOW 1.05 + .03. VOLTS DISPLAY SHOWS 30.5 + .3. POWER SOURCE LOAD DISPLAY SELECTOR SWITCH SELECTS POWER SOURCE (TRU OR GENERATOR) TO BE LOADMONITORED ON POWER SOURCE LOAD DISPLAY.

EXAMPLE: -.30 = 30% OF RATED DISCHARGE CAPACITY CHARGING AT A RATE GREATER THAN (1.00) IS INDICATED BY A FLASHING PLUS SIGN (+) PRECEDING THE NUMBER EXAMPLE: +1.12 = 12% OVER RATED CHARGE CAPACITY POWER SOURCE LOAD DISPLAY LUMINOUS DIGITAL INDICATOR DISPLAYS LOAD (CURRENT) DEMAND ON SELECTED POWER SOURCE (TRU OR GENERATOR) AS A DECIMAL FRACTION OF ITS MAXIMUM RATED LOAD CAPACITY. EXAMPLES: MAXIMUM LOAD DEMAND IS EXPRESSED AS 1.00; 60% OF MAXIMUM LOAD IS EXPRESSED AS .60. OVERLOADING OF POWER SOURCE IS INDICATED BY FLASHING PLUS SIGN (+) PRECEDING NUMBER. EXAMPLE: + 1.15 = 15% OVERLOAD.

Figure 24-21. DC Power Monitor Panel

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DC POWER MONITOR SYSTEM

NOTES

The DC power system is monitored from a DC SYSTEM indicator panel located adjacent to the DC CONTROL panel on the overhead console (Figure 24-21). The panel monitors voltage and load at the batteries, generators, and TRUs. Four luminous digital displays are provided. A BUS VOLTS display shows voltage at any one of the main, secondary, and essential buses, as selected with an adjacent bus voltage display selector switch. A power source LOAD display indicates the existing electrical load demand of any one of the DC generators or TRUs, as selected with the adjacent power source load display selector switch. The display is indicated in terms of a decimal fraction of the selected power source’s maximum rated load capacity (for the DC generators, this is equivalent to 300 amperes; for the TRUs, 200 amperes). For example, a reading of 1.00 indicates 100% load, a reading of 0.50 indicates 50% of maximum load, and a reading of 1.20 proceeded by a flashing “+’’ indicates a 20% overload. Battery load is shown by individual MAIN BATT and AUX BATT electrical LOAD displays that indicate existing battery charge or discharge rates in terms of decimal fractions of rated load (for both batteries that is equivalent to 100 amperes). A TEST button is provided for verification of all indicator displays and monitor circuitry. The verification is accomplished with a BITE (built-in test equipment) system.

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1

ENG 5 START (M5) 28V DC R ESS BUS

IGNITION OFF

NORM

2

1

4 HOLD

1

8

SELECT

3 2 5 6

START

SELECT

2

A

X1 X2

C1

A2

B2

5

1

2

B3 B1 A3

3

4

NORM

TD

5

ENG 2 IGN (L5)

5

G

MAN

1

OFF

ENGINE 2 IGNITION SWITCH 7411.S2

TIME DELAY RELAY K3

A1

D2

D1

OFF

ENGINE 1 IGNITION SWITCH 7411.S1 ENG 1 IGN (J5) 28V DC L ESS BUS

A

A

S2

A

PUSH BUTTON START

28V DC R ESS BUS

2

D2

C2

B2

X2

D3 D1 C3 C1 B3 B1

X1

START CONTROL RELAY K1

X1

D1

D3

C1

C3

B1

B3

TRIGGER SIGNAL

B2

C2

X2

D2

5 MAN

6

START CONTROL RELAY K2

4

NORM

3

B

B ENG 2 IGNITION EXCITER

A

ENG 1 IGNITION EXCITER

A

7411-P1

7411-PI

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

7

S1

START

SELECT

START

ENGINE START

INDENT CODE IS 8011-UNLESS OTHERWISE INDICATED

MANUAL

ENG 2

SELECT

ENG 1

1. 28V DC START MODE 0 VOLT GENERATING

ENG START SELECT

2 CONTACTOR CONTROL. SEE SHEET 2

NOTES:

3

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P 2

P 2

Revision 4

START/ TERMINATE

START CONTR

CURRENT LIMITING

SPEED SENSING SEE SHT. 2

SPEED SENSING SEE SHT. 2

CURRENT LIMITING

START CONTROL

START/TERMINATE

Figure 24-22. Engine Start Control—Electrical Schematic (Sheet 1 of 2)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ENGINE STARTING

b.

Starter-Generator Control— No. 2 Engine

c. With relay 8011-K2 energized, the following takes place:

NOTE The operation for starting the No. 2 engine is described. The procedure for starting the No. 1 engine first is similar except that the main bus tie must be closed prior to starting (Figure 24-22).

NOTE The starter-generator control sequence described is for the No. 2 engine with the No. 1 engine shut down. The sequence for the No. 1 engine is similar. To start the No. 2 engine, the No. 2 IGNITION switch is set to NORM. Selecting the start select switch to SELECT 2 initiates the following: 1. Start/terminate circuit for both generator control units (GCU-1 engine No. 1 and GCU-2 engine no. 2) is powered. 2 Holding coil energizes to retain the start select switch at SELECT 2. 3. Time delay relay 8011-K3 energizes to complete a current limiting circuit to GCU-1 and GCU-2 and to turn on the SELECT lens in the START switchlight.

Start control relay 8011-K2 and GCU-2 are energized.

(1) The START lens in the START switchlight illuminates. (2) Ignition circuits for No. 2 engine energize. (3) Start control circuit to GCU-2 is locked in. (4) Start/terminate control changes over to control by GCU-2. d. With GCU-2 energized, start contactor 2431-K2 energizes to connect the starter section of the No. 2 engine starter-generator to the right main feeder bus and GCU-2 to the 28-VDC right essential bus. T h e s t a r t e r- g e n e r a t o r, d r iv i n g through the turbomachinery accessory gearbox, rotates No. 2 engine high-pressure compressor. Fuel is switched on at a predetermined starter-generator speed, and light-up takes place. Starter-generator speed is sensed by the speed sensor (magnetic pickup) in the starter-generator, which relays speed signals to GCU-2.

For a detailed schematic analysis of engine start, refer to MSM Chapter 24.

4. START switchlight and start relay 8011-K2 are armed. 5.

With the SELECT lens on, pressing the START lens on the START switchlight initiates the following automatic sequence: a.

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Start control circuit GCU-2 is energized.

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ENGINE START IGNITION

1

OFF

START

2

SELECT

SELECT 1

2

28V DC L MAIN FEEDER BUS

2431-CB9

NORM START

28V DC L ESS BUS

MANUAL

12 X1

X2 11

R

P

G

P/O GENERATOR CONTROL UNIT (GCU -1)

2431-R1 ENGINE SPEED SENSING

CONTACTOR CONTROL

P/O 2431-P16

E

P/O 2431-P16

E K

START CONTROL SEE SHT.1

P

A3

ARM

P/O GENERATOR CONTROL UNIT (GCU -2)

TO NO. 2 GEN CAUTION LIGHT LOGIC CIRCUITS (CHAPTER 24)

P/O 2431-P12

B

H

B

H

ENGINE SPEED SENSING

SHUNT STARTERGENERATOR ENG NO. 1

N

K

G

13

TO NO. 1 GEN CAUTION LIGHT LOGIC CIRCUITS (CHAPTER 24)

P/O 2431-P11 S

A1 START CONTACTOR 2431-K1

P/O 2431-P11

START CONTROL SEE SHT.1

14

STARTERGENERATOR ENG NO. 2 ARM

DC CONTACTOR BOX

DC CONTACTOR BOX 2431-R2 SHUNT

S N

CONTACTOR CONTROL

R P/O 2431-P12 11 X1

13

A2

X2

START CONTACTOR 2431-K2

2431-CB23

NOTES : SEE SHEET 1

26V DC R ESS BUS

12

14

A1 28V DC R MAIN FEEDER BUS

Figure 24-22. Engine Start Control—Electrical Schematic (Sheet 2 of 2)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

6. At a predetermined starter-generator speed, the GCU-2 start control, start/terminate circuits, and start contactor 2431-K2 deenergize to initiate the following: a. With start control circuits deenergized, start relay 8011-K2 deenergizes to in turn deenergize ignition circuits, which completes the start circuit to both GCUs and turns off t h e S TA RT l e n s i n t h e S TA RT switchlight. b. With start/terminate circuits deenergized, the holding coil in the start select switch deenergizes, and the switch toggle snaps to the center off position. Time delay relay 8011K3 starts to time out. After two seconds delay, the relay operates to turn off the SELECT lens and deenergize current-limiting circuits for both GCUs. c.

With the start contactor 2431-K2 deenergized, the start section of the starter-generator is shut down, the starter-generator driven by the No. 2 engine high-pressure compressor through the turbomachinery accessory gearbox comes on line (provided the generator control switch is on), and the #2 DC GEN caution light extinguishes.

In this mode, the start procedure is the same as that described above except that the battery is charged by the output of the No. 2 engine starter-generator. The output of the on-line generator is limited by current-limiting circuits in GCU-1. Start mode off for engine cross-starting is similar to that described above for battery-only starting, with the exception that both startergenerators operate to supply power to the DC system.

Engine Starting—External Power Connected With the external power connected, 28-VDC power is applied to all buses through the left and right main feeder buses. In this mode, external power is the sole source of power, and the batteries and generators are disconnected from the buses. The start procedure for both engines is the same as that described for battery-only starting. The batteries and generators power the buses when the external power is switched off or removed.

NOTES

Cross-Starting NOTE Cross-starting the No. 1 engine with the No. 2 engine starter-generator on line is described. Crossstarting the No. 2 engine with the No. 1 engine starter-generator on line is similar. Cross-starting is accomplished with 28-VDC battery power applied to the right main feeder bus and the No. 2 engine generator on line.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Figure 24-23. AC Generator LEFT THREE-PHASE AC GENERATOR/EXTERNAL POWER INPUT AØ BØ CØ

LEFT

RIGHT THREE-PHASE AC GENERATOR/EXTERNAL POWER INPUT AØ BØ CØ 115-VAC VARIABLE-FREQUENCY CIRCUIT-BREAKER PANEL

RIGHT

BUS

L PROP DEICE PH A

L PROP DEICE PH B

ELEV HORN HT 1

STALL XDCR HTR 1

INTK LIP HTR ENG 1

L WDO HT

LEFT AC PH A

LEFT AC PH B

STBY HYD PMP 1 L TRU FUEL AUX PMP 1

L WSHLD HT

BUS

R PROP DEICE PH A

R PROP DEICE PH B

ELEV HORN HT 2

STALL XDCR HTR 2

R WSHLD HT

INTK LIP HTR ENG 2 LEFT AC PH C

RIGHT AC PH A

RIGHT AC PH B

RIGHT AC PH A

STBY HYD PMP 2 R TRU FUEL AUX PMP 2

Figure 24-24. Variable-Frequency AC Power Distribution

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AC POWER SYSTEM VARIABLE-FREQUENCY AC POWER Variable-frequency AC power is supplied by two AC generators, one mounted on the propeller gearbox of each engine. Generator output frequency varies from 333 to 528 Hz, depending on propeller speed; however, voltage remains at a constant 115 VAC (Figure 24-23). The output of each generator is normally applied through bus contactors to the related left and right variable-frequency AC buses. These buses are integral with the variable-frequency circuit-breaker panel in the flight compartment. Variable-frequency power distribution is shown in Figure 24-24. The generator output power is 115/200 volts, three-phase, variable-frequency, with a normal capacity of 20 KVA and an overload capacity of 30 KVA. The generator is a two-stage brushless,two-bearing design, with stationary and rotary excitation.

Lubrication and cooling of the generator is provided by internal oil circulation. Oil supplied from the engine enters an inlet port at the flange (mounting) edge under pressure. The oil is ducted to and through the hollow rotor shaft of the generator, exiting through eight spray nozzles which direct the flow into the rotor and stator windings. Bearing lubrication is provided by a controlled leakage device within the rotor shaft. The oil flows through a screen to a collector sump in the bottom of the generator and is returned to the engine through the outlet port by action of the oil scavenge pump. Each generator is governed by a generator control unit (GCU), which monitors voltage, current, and speed to automatically initiate and regulate generator output at a constant 115 ±2.8 volts per phase throughout the normal speed range. The GCUs also provide automatic overvoltage, undervoltage, underfrequency, and bus fault protection to each generator.

NOTES

The first-stage generator, mounted on the rotor shaft, is excited by the stationary field. This generates an AC current which is rectified by diodes mounted on the shaft to supply current to the main field for excitation of the secondor main-stage stator output windings. A magnetic speed sensor and current transformer are also mounted in the generator housing. An overtemperature switch is installed in the generator. When the stator windings reach a temperature of 177° C (350.6° F), the switch contacts close to operate the AC GEN HOT lights on the caution panel. The switch automatically resets when the temperature returns to a safe level. The output frequency of the generator varies with propeller speed and ranges from 333 Hz at 10,000 rpm to 528 Hz at 15,850 rpm; however, the output voltage is controlled and regulated by the GCU.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

REMOTE C/B

MODULE-BLOCK

RELAY

CURRENT TRANSFORMER

GCU

VOLTAGE SENSOR

COVER ASSEMBLY— FUSE LIMITERS

TPYICAL AC CONTACTOR BOX (RH NACELLE) NOTE: LH NACELLE SIMILAR

Figure 24-25. AC Contactor Box and Components

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

AC Contactor Box There are two AC contactor boxes (Figure 2425), one mounted in each engine nacelle, containing most of the components for the left and right AC systems. Each box contains a bus contactor, an external power contactor, a generator control unit, current transformers, output feeder fuses, and control circuit breakers for the associated side. Also contained in each box is the remote-control circuit breaker for the left or right standby hydraulic pump. The right AC contactor box holds the external power protection unit and fuses. All electrical connections are via multipin connectors mounted on two sides of the box assemblies. The output feeders consist of three No. 12 wires, routed from each contactor box to the left and right AC variable-frequency buses circuitbreaker panel in the flight compartment.

Contactors Two bus contactors (left and right) switch the three-phase power from the generator to the associated bus. If power from the associated generator is not available, the bus contactor allows that bus to be powered from the opposite generator (crosstie position). The crosstie position is also used when external power is selected.

GCU of the inoperative generator switches the associated bus contactor to the crosstie position. The operational generator then supplies both the left and right variable-frequency buses.

Current Transformers Six current transformers are installed in the system (T1 through T6). Transformers T5 and T6 measure total generator output current for the AC power monitor system. Transformers T1 and T4 measure total current flow in the main and crosstie feeders. The sensed output is used in the GCU differential current-sensing circuits. Transformers T2 and T3 measure crosstie current. They sense current flow only when the bus contactors are in the crosstie position. Their output is summed with the output of the opposite T1 or T4 for differential current sensing within the GCU. Each current transformer assembly consists of three single-phase toroidal current transformers, installed in a lightweight aluminum housing. Each line of the three-phase system to be monitored is passed through its toroid. Electrical connections to the toroids are via a multipin connector. The mounting holes for the assembly are asymmetrical to avoid improper installation.

The left and right bus contactors are hermetically sealed, three-position center-off, with a three-pole double-throw (3PDT) contact arrangement. Separate coil windings, designated X and Y, are used to energize the contact arm to either of the two closed positions. Heavy-threaded studs provide electrical connections to external circuits. An automatic crosstie function, controlled by the GCU logic circuits, ensures that all variablefrequency buses are powered when only one AC generator is on line, such as during engine start or following in-flight failure or shutdown. Whenever such a fault condition exists, the

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CURR XMFR

CURR INDIC

CURR XMFR

DIFF CURRENT SENSING NO. 1 AC GEN CONT UNIT (AC GCU) DIFF CURRENT SENSING PROT L GEN ON LINE R GEN ON LINE

NO.1 AC GEN HOT

NO.2 AC GEN HOT

L AC BUS

R AC BUS

NO. 1 AC NO. 1 AC NO.1 AC GEN CONT GEN CONT NO.2 AC GEN GEN SWITCH SWITCH

GENERATOR INHIBIT FROM L ESS BUS (28 VDC)

L BUS CONTACTOR

R STBY HYD PUMP

R EXT PWR CONT

EXT AC PWR ON

AØ

CURR XMFR

CURR XMFR

FROM R ESS BUS (28 VDC)

L EXT PWR CONT

R ESS BUS (28 VDC)

DIFF CURRENT SENSING CURR NO. 2 AC INDIC GEN CONT UNIT (AC GCU) DIFF CURRENT SENSING PROT L GEN ON LINE R GEN ON LINE

AC EXT PWR CONT SW

R BUS CONTACTOR

L STBY HYD PUMP

AC EXTERNAL POWER PROTECTION UNIT

AØ

BØ

BØ L TRU

R TRU

CØ

CØ

L VAR FREQ AC BUSES

AC EXTERNAL POWER RECEPTACLE

R VAR FREQ AC BUSES

Figure 24-26. AC Variable Frequency Generation—Simplified

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AC GENERATOR CONTROL UNIT (GCU) The GCU is all solid-state, with the exception of two control relays which are hermetically sealed. The components and printed circuit boards are encased in a single-piece aluminum cover on which is mounted a multipin connector. Each unit is located in an AC contactor box in each engine nacelle. The GCU contains the voltage regulator and control circuits for the generator, protection circuits, and bus contactor control. Contactor control includes the automatic switchover function of the feeders and buses in the event of generator failure. With both generators turned on, each GCU monitors the following functions for control of the generator and contactor: • Generator output voltage on each phase measured at the bus contactors and supplied to pins F, E, and D through circuit breakers 3, 5, and 7 (left) and 4, 6, and 8 (right) located in the AC contactor boxes • Feeder current in each phase, developed in transformers T1 and T4 and supplied to pins M, N, and P for differential current sensing. • Internal generator current, developed from current transformers within the generators and supplied to pins R, S, and T • Generator speed (rpm), developed from the magnetic speed sensor within the generator and supplied to pins U and V • Operational status information from the other GCU. This information is provided by direct interconnections between the GCUs on pins m, t, n, b, s, i, j, and r.

plied by the essential DC buses, the essential buses must be energized for the AC generators to function. However, crossconnection circuitry of the GCU power supplies ensures that a GCU losing its essential bus feed can draw DC power from the opposite GCU. Manual activation and deactivation of the AC generators is controlled by GEN 1 and GEN 2 control switches on the AC CONTROL panel in the flight compartment (Figure 24-27). When switched on, each generator control switch signals its GCU to activate the generator and connect it to the bus when generator voltage rises to a minimum value. During initial startup, the GCU, once switched on, connects the generator to its bus as soon as voltage is above 42 volts and automatically shuts down and removes the generator from the bus if output voltage exceeds 125. The generator control switches also have a GCU reset function with which an AC generator that has been automatically shut down may be reactivated by off–on selection of its control switch. The AC generators are protected from bus faults by the GCUs, which monitor generator output voltage at the bus contactors to detect the excessive voltage drop that results from a short circuit on a bus. Once such a drop below 90 volts is detected, the GCU isolates the bus by opening its bus contactor and illuminates the appropriate L AC BUS or R AC BUS caution lights on the caution light panel (Appendix B). The #1 AC GEN and #2 AC GEN caution lights illuminate in response to GCU signals and indicate that the affected generator is shut down. Caution lights are also provided to warn of generator overheat. These #1 AC GEN HOT and #2 AC GEN HOT lights illuminate when the related AC generator temperature exceeds a predetermined limit. The caution lights extinguish when the related generator cools below the overheat point.

Internal logic circuits in each GCU control the operation of bus contactors that connect the AC generators to the variable-frequency buses. Since the GCUs operate on DC power sup-

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AC GENERATOR CONTROL SWITCH AT GEN 1 OR 2 POSITION, THE ASSOCIATED GCU IS ARMED TO ACTIVATE ITS GENERATOR WHEN AT OPERATIONAL SPEED. THE OFF POSITION SHUTS DOWN THE GENERATOR, OR FOLLOWING SHUTDOWN BY EITHER MANUAL OR AUTOMATIC MEANS, RESETS THE GCU FOR SUBSEQUENT REACTIVATION. INVERTER CONTROL SWITCHES THREE SWITCHES CONTROL ACTIVATION OF THEIR RESPECTIVE INVERTERS.

EXTERNAL POWER ADVISORY LIGHT (GREEN)

PRIMARY SWITCH, WHEN POSITIONED TO PRIMARY, ENERGIZES THE PRIMARY INVERTERS TO POWER THE LEFT 115-VAC, 400-HZ BUS.

ILLUMINATES WHEN EXTERNAL POWER IS AVAILABLE TO THE VARIABLE-FREQUENCY AC BUSES EXTERNAL POWER SWITCH WHEN POSITIONED TO EXT PWR, AN AC EXTERNAL POWER SOURCE IS CONNECTED TO THE VARIABLE-FREQUENCY AC BUSES, AND THE AC GENERATORS ARE ISOLATED FROM THE BUSES.

AUXILIARY SWITCH, WHEN POSITIONED TO L OR R, ENERGIZES THE AUXILIARY INVERTER AND APPLIES IT TO THE LEFT OR RIGHT 115-VAC, 400-HZ BUS. SECONDARY SWITCH, WHEN POSITIONED TO SECONDARY, ENERGIZES THE SECONDARY INVERTER TO POWER THE RIGHT 115-VAC, 400-HZ BUS.

Figure 24-27. AC CONTROL Panel

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AC Power Monitoring System The AC power monitor panel (Figure 24-32), labeled “AC SYSTEM,’’ is located adjacent to the AC control panel. Located on the right side of the panel, above the label “VARIABLE FREQUENCY,’’ are two digital readout windows labeled “VOLTS’’ and “LOAD’’ and a six-position rotary selector switch. The digital readouts indicate voltage and current readings on the selected left or right phase. The LOAD display window monitors load currents expressed as a percentage factor of the rated circuit load. For example, a readout of 1.00 indicates full load (87 amperes per phase),a readout of .50 indicates half load, and a readout of 1.20 indicates a 20% overload. An overload condition is advised by a flashing positive sign (+) preceding the readout. Pressing the TEST pushbutton tests the monitor display and circuit operation by using a BITE (built-in test equipment) system. During the test all segments of both the VOLTS and the LOAD digital display characters illuminate. When held longer than approximately two seconds, the internal test circuits simulate a voltage of 150 ±3 and loads of 1.05 ±.03, which are displayed on both sections of the panel.

tioning of the bus contactors via the GCUs and connect the receptacle directly to the left and right variable-frequency buses. An advisory light on the AC CONTROL panel illum i n a t e s t o s h ow t h a t e x t e r n a l p ow e r i s available to the buses. While external power is applied, the AC generators are isolated from the variable-frequency buses. An external power protection unit monitors the external power source for faults. If a malfunction is detected (abnormal voltage, current, or phase rotation), the unit switches the external power system off.

NOTES

Voltage readings are from individual circuit breakers on each phase of the left and right power buses; current indications are obtained from transformer T5 on the left generator output and T6 on the right generator output.

AC External Power Operation of the variable-frequency AC system from a ground power source is provided for through an external power receptacle and control circuitry installed in the inboard side of the right nacelle, to the rear of the main shock strut hinge point. Application of external power to the variablefrequency buses is controlled by the EXT POWER switch on the AC CONTROL panel. When external power is selected, contactors operate in conjunction with crosstie posi-

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{

AC POWER MONITOR AC CONTROL PANEL

400-HZ AUTOTRANSFORMERS (REAR OF BULKHEAD X182.00)

PRIMARY INVERTER

SECONDARY INVERTER

PARALLELING CONTROL BOX 400-HZ CONTROL BOX AUXILIARY INVERTER

INVERTER WARNING CONTROL BOX

Figure 24-28. AC 400-Hz Power System Component Locations

INVERTER CONTROL SWITCHES AC CONTROL

THREE SWITCHES CONTROL ACTIVATION OF THEIR RESPECTIVE INVERTERS. OFF

GEN 1

PRIMARY SWITCH ENERGIZES PRIMARY INVERTER TO

GEN 2

INVERTERS

POWER LEFT 115-VOLT, 400 HZ BUS WHEN SELECTED TO PRIMARY.

SECONDARY SWITCH ENERGIZES SECONDARY OFF EXT POWER

PRIMARY

L

R

OFF

OFF SECONDARY AUXILIARY

INVERTER TO POWER RIGHT 115-VOLT , 400 HZ BUS WHEN SELECTED TO SECONDARY.

AUXILIARY SWITCH (THREE POSITIONS) ENERGIZES AUXILIARY INVERTER AND APPLIES IT TO EITHER LEFT OR RIGHT 115-VOLT, 400 HZ BUS WHEN SELECTED TO L OR R AS APPROPRIATE.

Figure 24-29. Fixed Frequency

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FIXED-FREQUENCY AC POWER Fixed-frequency AC power is supplied at a constant 115 volts, 400 Hz from three solidstate static inverters installed in the flight compartment (Figure 24-29). The inverters are powered by 28 VDC supplied via the left essential (primary inverter), right essential (secondary inverter), and left main (auxiliary inverter) DC buses. A schematic of the fixedfrequency system is shown in Figure 24-30; fixed-frequency power distribution is shown in Figure 24-31.

faulted bus and its inverter(s). The inverter(s) connected to the faulted bus may subsequently fail if the overload persists.

NOTES

Inverter output is applied to left and right 115-volt, 400-Hz buses for the operation of 115-volt avionics equipment. Two step-down transformers, one powered from each 115volt fixed-frequency bus, reduce part of the fixed-frequency supply to 26 VAC. Transformer output is applied to left and right 26vo l t , 4 0 0 - H z bu s e s f o r t h e o p e r a t i o n o f 26-volt avionics equipment. Both the 115volt and 26-volt fixed-frequency buses are integral with the avionics circuit-breaker panel in the flight compartment. A paralleling control box, operating in conjunction with inverter control switches (Figure 24-29) in the flight compartment, controls the application of the three inverters to the 115volt buses. The box applies primary inverter output to the left bus and secondary inverter output to the right bus and allows selection of the auxiliary inverter to either bus. The buses are tied, via a bus-tie circuit breaker, so demands on all buses are shared by the operational inverters. Load-sharing circuits within the paralleling control box balance the load demands among the inverters to maintain uniform output from each. The circuit breaker in the bus-tie circuit provides a measure of protection from bus faults (short circuits). A bus fault resulting in abnormally high current flow (greater than a level equivalent to one inverter’s maximum output) across the bus tie opens the circuit breaker, isolating the faulted bus from the un-

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L ESS BUS

L MAIN FEEDER BUS

R ESS BUS

PRI INV

AUX INV

SEC INV

PRI INV SW

PRI INV

AUX INV SW

AUX INV

SEC INV SW

SEC INV

PARALLELING CONTROL BOX R 115-VAC BUS

L 115-VAC BUS 26-VAC XFMR

26-VAC XFMR

L 26-VAC BUS

R 26-VAC BUS

L 26 AC

R 26 AC

LEGEND 28-VDC POWER INVERTER POWER TRANSFORMER POWER

Figure 24-30. Fixed-Frequency AC Power Schematic AUXILIARY INVERTER INPUT (R SELECTION)

AUXILIARY INVERTER INPUT (L SELECTION)

PRIMARY INVERTER INPUT

LEFT 26-VAC BUS VOR 1 ADF 1 HDG 1 ATT 1 ALT 1 HDG CRS ERROR 1 SYM GEN 1 FLAP POSN 1 HYD QTY 1 L 26V FAIL SURF POSN IND

SECONDARY INVERTER INPUT

RIGHT 115V 26-VAC BUS AUX AUX BUS TIE LEFT INV IN INV IN VOR 2 115-VAC ATT PNL LTG NAV SW 2 ADF 2 BUS NAV SW 1 FGC 2 HDG 2 FGC 1 SYM GEN 2 ATT 2 SYM GEN 1 WEA RDR ALT 2 GPWS CVR HDG CRS ERROR 2 RIGHT FDR 115/26-VAC SYM GEN 2 115-VAC 115/26-VAC XFMR LT FLAP POSN 2 BUS XFMR LT HYD QTY 2 R 26V FAIL AVIONICS CIRCUIT-BREAKER PANEL L 26-VAC TRANSFORMER R 26-VAC TRANSFORMER

Figure 24-31. Fixed-Frequency AC Power Distribution

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Inverter activation and deactivation are controlled by three INVERTERS switches on the AC CONTROL panel in the flight compartment (Figure 24-29). The PRIMARY and SECONDARY switches are two-position, on-off switches. The AUXILIARY switch has three positions, which provide for the activation and application of the auxiliary inverter to either the left or right fixed-frequency buses, as desired. The paralleling control box is mounted on the center shelf of the left circuit-breaker (Figure 24-28) console, adjacent to the secondary inverter.

NOTES

NOTE A failed inverter as indicated by a MASTER CAUTION light must be turned off manually by the related inverter control switch located on the AC CONTROL panel.

An inverter warning control unit monitors the three inverters for output and controls the illumination of the PRI INV, SEC INV, and AUX INV lights located on the caution light panel (Appendix B). Each caution light illuminates when the warning control unit senses that the related inverter is off or has failed. The inverter warning control box is mounted on top of the 400-Hz control box, located on the bottom shelf of the left circuit-breaker console. Two 26-VAC, 400-Hz autotransformers and power correction capacitors are installed on the rear face of the bulkhead at FS X182.00, behind the pilot’s circuit-breaker console. The transformers are supplied from the left and right 115-VAC, 400-Hz buses through the 115/26-VAC XFMR LT circuit breaker (CB13) and the 115/26-VAC XFMR RT circuit breaker (CB12). The outputs are connected to the related 26VAC, 400-Hz bus. L 26 AC and R 26 AC caution lights are also provided to warn of power loss at the 26-volt fixed-frequency buses. Component locations are shown in Figure 24-28.

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INVERTER VOLTAGE DISPLAY DISPLAY VOLTAGE OUTPUT OF SELECTED INVERTER INVERTER LOAD DISPLAY DISPLAY LOAD (CURRENT) ON SELECTED INVERTER AS A DECIMAL FRACTION OF ITS MAXIMUM LOAD CAPACITY. EXAMPLES: MAXIMUM LOAD IS EXPRESSED AS 1.00; 60% OF MAXIMUM LOAD IS EXPRESSED AS .60. OVERLOADING OF INVERTER IS INDICATED BY A FLASHING PLUS SIGN (+) PRECEDING NUMBER. EXAMPLE: +1.05 =5% OVERLOAD

VARIABLE-FREQUENCY LOAD DISPLAY INVERTER SELECTOR SELECTS THE INVERTER TO BE MONITORED AC MONITOR SYSTEM TEST BUTTON WHEN PRESSED AND HELD, ALL VOLTS AND LOAD DISPLAY SEGMENTS ILLUMINATE FOR TWO SECONDS, FOLLOWED BY APPEARANCE OF THE FOLLOWING TEST VALUES: LOAD DISPLAYS SHOWS 1.05 ±.03; VOLTS DISPLAYS SHOWS 150 ±3.

DISPLAYS LOAD (CURRENT) AT ANY ONE PHASE OF EITHER AC GENERATOR'S OUTPUT SELECTED ON PHASE SELECTOR AS A DECIMAL FRACTION OF MAXIMUM LOAD CAPACITY. VARIABLE-FREQUENCY VOLTAGE DISPLAY DISPLAYS VOLTAGE AT ANY ONE PHASE OF EITHER AC GENERATOR'S OUTPUT AS SELECTED ON PHASE MONITORING SELECTOR VARIABLE-FREQUENCY PHASE SELECTOR SELECTS ANY ONE OF THE SIX VARIABLE-FREQUENCY AC PHASES TO BE MONITORED

Figure 24-32. AC Power Monitor

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AC POWER MONITOR

NOTES

AC power is monitored from an AC SYSTEM indicator panel (Figure 24-32), located adjacent to the AC CONTROL panel. The AC SYSTEM indicator panel receives output voltage and load information from the two AC generators and three inverters via current transformers at each power source. The left side of the panel contains two luminous digital INVERTERS VOLTS and LOAD indicators. The indicators display output voltage and load for any one of the three static inverters as selected with the three-position INVERTERS switch. Inverter load is presented as a decimal fraction of maximum inverter capacity. For example, a readout of 1.00 indicates full load (3.5 amperes per phase), a readout of 0.50 indicates half load, and a readout of 1.20 indicates a 20% overload. An overload condition is advised by a flashing positive sign (+) preceding the readout. The right side of the panel contains two luminous digital VARIABLE FREQUENCY VOLTS and LOAD indicators, which were covered in the Variable-Frequency AC Power section of this chapter. A TEST button is provided for verification of all indicator displays and monitor circuitry. When the button is pressed, a BITE circuit produces a predetermined display indication and thereby verifies proper operation.

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ELECTRICAL SYSTEM CIRCUIT PROTECTION Circuit protection is provided for electrical system power sources, component control circuits, and bus distribution by circuit breakers located on the four flight compartment circuitbreaker panels and the DC and AC contactor boxes in the nose compartment and right and left nacelles. Those breakers which are located on the flight compartment panels are illustrated in Figures 24-33, 24-34, and 24-35.

Figure 24-33. Right DC Circuit-Breaker Panel

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Figure 24-34. Left DC Circuit-Breaker Panel

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AC VARIABLE FREQUENCY CIRCUIT-BREAKER PANEL

AVIONICS CIRCUIT-BREAKER PANEL

Figure 24-35. Varible Frequency and Avionics Circuit-Breaker Panels

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Figure 24-36. Typical DC Power Distribution

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MAINTENANCE CONSIDERATIONS REMOVAL/INSTALLATION

5.

After installation, reconnect the batteries and the external power source as applicable. Remove the clips and push in the circuit breakers; then perform an operational test on the unit.

BATTERY

General Unless required by applicable maintenance practices, removal and installation procedures for electrical equipment are considered selfevident. Equipment grounding and bonding surfaces must be cleaned to ensure good conductivity, and after installation any bared surfaces must be refinished.

Removal and installation of the auxiliary battery is similar to that of the main battery. However, if the auxiliary battery is to be removed, it is necessary to first remove the main battery in order to facilitate access to the auxiliary.

WARNING

Safety Precautions WARNING Failure to observe safety precautions before carrying out maintenance practices on or near electrically controlled or operated equipment could result in injury to personnel and/or damage to equipment.

1.

Ensure that the airplane is properly grounded.

2.

Ensure that the BATTERY MASTER, MAIN BATT, AUX BATT, and EXT PWR switches on the DC control panel and the EXT POWER switch on the AC control panel are in OFF.

3.

Disconnect the left and right batteries and the external power sources. Pull out the applicable system circuit breakers and insert clips to prevent push-to-reset.

4.

If units removed are not to be immediately replaced, ensure that the connector ends are capped and stowed and that all loose wires are insulated and stowed.

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Battery electrolyte consists of a caustic solution of potassium hydroxide which is hazardous to personnel. Avoid contact with skin and eyes. If skin contact occurs, bathe the affected area with large quantities of water and neutralize with boric acid solution or vinegar. If eye contact occurs, flush with water and obtain immediate medical attention.

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CAUTION

NOTES

Ensure that the BATTERY MASTER switch is OFF and that all external power sources are removed from the airplane before disconnecting or removing the battery.

CAUTION If external power is to be applied to the airplane with the battery removed, ensure that the battery connector is protected from accidental contact with the airframe or components.

CAUTION The battery is extremely heavy (approximately 80 pounds) and must be carefully handled, preferably by two persons, to avoid damage to the battery or injury to individuals.

CAUTION To avoid unnecessary loss of battery power, switch off all electrical equipment not specifically required for testing.

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CHAPTER 26 FIRE PROTECTION CONTENTS Page INTRODUCTION ................................................................................................................. 26-1 GENERAL ............................................................................................................................ 26-3 ENGINE FIRE PROTECTION SYSTEM ............................................................................ 26-5 Engine Fire Detection .................................................................................................... 26-5 Components Description and Operation........................................................................ 26-5 Fire Detection—Series 300 ......................................................................................... 26-17 Engine Fire Extinguishing ........................................................................................... 26-19 LAVATORY FIRE PROTECTION SYSTEM..................................................................... 26-25 Lavatory Smoke Detection—Mod 8/0266................................................................... 26-25 Lavatory Smoke Detection—Series 300 ..................................................................... 26-27 Lavatory Fire Extinguishing ........................................................................................ 26-29 BAGGAGE COMPARTMENT SMOKE DETECTION SYSTEM—SERIES 100/300 .... 26-31 General......................................................................................................................... 26-31 System Operation ........................................................................................................ 26-31 System Test.................................................................................................................. 26-31 PORTABLE FIRE EXTINGUISHERS—SERIES 100 ...................................................... 26-33 PORTABLE FIRE EXTINGUISHERS—SERIES 300 ...................................................... 26-33 MAINTENANCE CONSIDERATIONS ............................................................................ 26-34 Maintenance Practices ................................................................................................. 26-34 Servicing...................................................................................................................... 26-35 Inspections................................................................................................................... 26-35

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FAULT ANALYSIS............................................................................................................. 26-35 FUNCTIONAL CHECKS................................................................................................... 26-35

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ILLUSTRATIONS Figure

Title

Page

26-1

Fire Protection Panel and Warning Lights ............................................................. 26-2

26-2

Smoke Detection Indication and Test..................................................................... 26-2

26-3

Fire Detection Sensor Loops Installation............................................................... 26-4

26-4

Fire Detector Operating Principle (Physical)......................................................... 26-6

26-5

Fire Detector Operating Principle (Functional) ..................................................... 26-6

26-6

Fire Detection Control Units.................................................................................. 26-8

26-7

Fire Protection Panel and Warning Lights ........................................................... 26-10

26-8

Glareshield and Master Warning Lights .............................................................. 26-10

26-9

Fire Detection System Electrical Schematic ....................................................... 26-12

26-10

Fire Detection (300 Series) (Mod 8/1835)........................................................... 26-16

26-11

Fire Protection Panel............................................................................................ 26-18

26-12

Fire Extinguisher Bottles Installation .................................................................. 26-20

26-13

Distribution Lines ................................................................................................ 26-20

26-14

Fire Extinguisher Discharge Indicator Discs ....................................................... 26-21

26-15

Engine Fire-Extinguishing System ...................................................................... 26-22

26-16

Lavatory Smoke Detection System and Electrical Schematic ............................. 26-24

26-17

Lavatory Smoke Detection System—Series 300 ................................................. 26-26

27-18

Lavatory Smoke Detection Electrical Schematic—Series 300............................ 26-28

26-19

Lavatory Fire Extinguisher .................................................................................. 26-28

26-20

Baggage Compartment Smoke Detection System—Series 100 and 300............. 26-30

26-21

Baggage Compartment Smoke Detection System Electrical Schematic ............. 26-30

26-22

Portable Fire Extinguishers—Series 100 ............................................................. 26-32

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CHAPTER 26 FIRE PROTECTION FIRE WARN

FIRE PULL

INTRODUCTION Fire protection in the Dash 8 consists of fire detection, smoke detection, and fire-extinguishing systems. Fire detection in the rear baggage compartment and in the lavatory is provided by smoke detectors. Fire detection in the nacelles is provided by dual sensor loops. Fire extinguishing consists of an independent two-bottle engine fire-extinguishing system, an automatic single-bottle lavatory fire-extinguishing system, three hand-operated portable fire extinguishers, and an optional single-bottle APU fire-extinguishing system.

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CHECK FIRE DET WARNING LIGHT (RED)

WARNING PRESS TO RESET ENGINE FIRE PRESS TO RESET LEFT MASTER FIRE WARNING LIGHT (FLASHING)

ILLUMINATES FLASHING CONCURRENT WITH BOTH ALARM AND FAULT CONDITIONS. WILL RESET TO STEADY ON WARNING RESET.

ILLUMINATES WITH ANY WARNING LIGHT, FLASHING UNTIL RESET.

FAULT LIGHTS (4) ILLUMINATION OF ANY LENS (CONCURRENT WITH THE CHECK FIRE DET WARNING LIGHT) INDICATES A MALFUNCTION HAS BEEN DETECTED WITHIN THE ASSOCIATED DETECTOR LOOP CIRCUIT.

ILLUMINATES CONCURRENT WITH PULL FUEL OFF LIGHT AND CHECK FIRE DET WARNING LIGHT TO WARN OF A DETECTED FIRE IN EITHER NACELLE.

ENGINE FIRE PRESS TO RESET

EXTG AFT BTL

FUEL VALVE OPEN CLOSED

EXTG AFT BTL

FUEL VALVE OPEN CLOSED

RIGHT MASTER FIRE WARNING LIGHT (FLASHING) ILLUMINATES CONCURRENT WITH PULL FUEL OFF LIGHT AND CHECK FIRE DET WARNING LIGHT TO WARN OF A DETECTED FIRE IN EITHER NACELLE

BAGGAGE SMOKE WARNING FWD BTL

FWD BTL A

BOTH

B

A

BOTH

TEST 1

B OFF

FAULT A

FAULT B

FAULT A

LOOP SELECTION

TEST 2

LOOP SELECTION

ENGINE 1

PULL FUEL OFF

FAULT B

FUEL EMERGENCY SHUTOFF VALVE LIGHTS ILLUMINATES TO INDICATE VALVE OPEN OR CLOSED

ENGINE 2 TEST DETECTION

PULL FUEL OFF

DETECTION SYSTEM TEST SWITCH THREE - POSITION CENTER - OFF SWITCH SIMULATES ALARM AND FAULT CONDITIONS IN THE ASSOCIATED NACELLE'S FIRE DETECTION SYSTEM WHEN SELECTED AND HELD AT ITS MOMENTARY LEFT (ENGINE 1 ) OR RIGHT (ENGINE 2 ) POSITION.

PULL FUEL OFF HANDLE LIGHT (2)

LOOP ARMING SELECTOR SWITCH (2) NORMALLY LEFT AT BOTH (BOTH LOOPS ARMED). A AND B POSITIONS ARE FOR DISARMING INDIVIDUAL MALFUNCTIONING LOOPS, I.E. SELECTING A DISARMS LOOP B SELECTING B DISARMS LOOP A.

ILLUMINATION OF THE ENGINE FIRE, PULL FUEL OFF, AND FAULT LIGHTS OF THE AFFECTED SIDE PLUS THE CHECK FIRE DET WARNING LIGHT INDICATES SYSTEM SERVICEABILITY. NOTE: DISARMED LOOPS ARE EXCLUDED FROM THE FAULT TEST FUNCTION.

ILLUMINATES CONCURRENT WITH ENGINE FIRE LIGHT AND CHECK FIRE DET WARNING LIGHT UPON DETECTION OF FIRE BY EITHER LOOP IN THE ASSOCIATED NACELLE.

Figure 26-1. Fire Protection Panel and Warning Lights SMOKE DETECTOR TEST SWITCH ACTIVATES ALARM CIRCUIT OF RELATED NO. 1 OR NO. 2 SMOKE DETECTOR WHEN HELD AT TEST 1OR TEST 2 APPROPRIATELY. ILLUMINATION OF SMOKE WARNING LIGHT DURING TEST INDICATES CIRCUIT SERVICEABILITY.

FUEL VALVE CLOSED

SMOKE DETECTOR TEST SWITCHES (S.O.O. only) ACTIVATES ALARM CIRCUIT OF RELATED NO. 1 OR NO. 2 SMOKE DETECTOR AND , IF INSTALLED, NO. 3 OR NO. 4 SMOKE DETECTOR WHEN HELD AT TEST 1OR TEST 2 AND TEST 3 OR 4 APPROPRIATELY. ILLUMINATION OF SMOKE WARNING LIGHT DURING TEST INDICATES CIRCUIT SERVICEABILITY.

EXTG AFT BTL

FUEL VALVE OPEN CLOSED

BAGGAGE SMOKE WARNING FWD BTL BOTH

B

A

BOTH

BAGGAGE SMOKE WARNING GARGO

TEST 1

B

FAULT A

SELECTION

FAULT B

TEST 2

LOOP SELECTION

E1

F U L L

1 2

1 4

OFF TEST 2

TEST DETECTION

PULL FUEL OFF

FIRE PROTECTION PANEL

WARNING LIGHT

TEST 3 OFF

ENGINE 2 TEST 4

EL OFF

SMOKE

TEST 1

OFF

FAULT B

ILLUMINATES UPON DETECTION OF SMOKE BY ANY SMOKE DETECTOR, OR DURING TEST OF FUNCTIONING SMOKE DETECTOR CIRCUIT.

FIRE PROTECTION PANEL (S.O.O.)

Figure 26-2. Smoke Detection Indication and Test

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GENERAL

NOTES

Nacelle fire or overheat is detected by continuous-loop sensors routed around vital areas of the engines. Duplicate loops parallel the installation to provide additional protection. Nacelle fire extinguishing is provided by two extinguisher bottles with crossover capability to the opposite nacelle. Warning of smoke in the baggage compartment is provided by smoke detectors on the roof and aft face of the compartment divider bulkhead. Smoke detected by either unit illuminates a warning light in the flight compartment. A modification adds a smoke detection system to illuminate a warning light on the flight attendant’s panel and sound a chime when smoke is detected in the lavatory. An extinguisher automatically discharges into the disposal bin, as required. An optional APU fire protection system utilizes a continuous-loop sensor around vital areas of the APU to detect a fire or overheat condition. APU shutdown and extinguisher bottle discharge occur if a fire or overheat condition exists. APU fire protection and detection will be covered in Chapter 49 of this manual.

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ZONE 3A—LEADING EDGE ALONG SIDES AND TOP

MOD 8/0278 ZONE 1—COMBUSTION ALONG SIDE AND UNDER EXHAUST DUCT ZONE 2—FORWARD ALONG STRUCTURE

ZONE 3—WHEEL WELL ALONG UPPER SIDES

POST MOD 8/0278 ZONE 1—COMBUSTION ALONG SIDE AND OVER EXHAUST DUCT

STAINLESS STEEL TUBE

HELIUM

INERT METALLIC MATERIAL AND HYDROGEN

FIRE DETECTION SENSOR LOOPS ROUTING IN NACELLE LOOP A RESPONDER

TITANIUM WIRE

LOOP B RESPONDER

NORMALLY CLOSED INTEGRITY SWITCH

ALARM/ INTEGRITY SIGNAL NORMALLY OPEN ALARM SWITCH

RESISTOR

RESPONDERS

Figure 26-3. Fire Detection Sensor Loops Installation

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ENGINE FIRE PROTECTION SYSTEM ENGINE FIRE DETECTION General Two independent fire detection systems, one in each nacelle, provide warning if a fire or overheat condition occurs. Each system consists of two separate, parallel-routed fire detection sensor loops with responders, a fire detection control unit, control switches, and warning lights.

in the center of the tube. The space between the inert material and the tube wall is charged with helium gas. One end of the tube is sealed; the other end is fitted with a responder. The gases in the loop create and maintain a minimum pressure on the attached responder. Overall heating of the loop causes a pressure i n c r e a s e o n a s w i t c h i n t h e r e s p o n d e r. Localized heating of the loop causes a release of hydrogen gas from the inert metallic material, creating a similar effect by increasing pressure on the switch.

NOTES

The sensor loops, designated A and B, pass through fire zones in each nacelle and are connected to fire detection control units to form a fire detection system for each nacelle. Each loop is filled with gases that expand as the temperature increases. At a preset pressure, an alarm switch in the responder closes, triggering alarm circuits in the control unit, generating signals to illuminate warning lights on the fire protection and glareshield panels. The lights also illuminate if a malfunction occurs in a sensor circuit. The detection system uses 28 VDC, both systems being supplied by the left and right essential buses.

COMPONENTS DESCRIPTION AND OPERATION Sensor Loops Two separate fire detection sensor loops are installed in each nacelle to detect fire or overheat conditions. The loops are installed in parallel and are routed through designated zones (Figure 26-3). The loops are stainless steel capillary tubes .063 inch (1.6 mm) in diameter, with a wall thickness of .018 inch (.45 mm). A titanium wire surrounded by a spongy inert metallic material impregnated with hydrogen gas is

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ALARM SWITCH (DIAPHRAGM) (N.O.)

ELECTRICAL ISOLATOR/PNEUMATIC MANIFOLD

POWER

SENSOR TUBE AVERAGING GAS (HELIUM)

SIGNAL

CORE ELEMENT (METAL HYDRIDE)

INTEGRITY SWITCH (N.C.)

ELECTRICAL INTERFACE

DUAL RESPONDER

PNEUMATIC SENSOR

Figure 26-4. Fire Detector Operating Principle (Physical)

POWER SIGNAL

OVERHEAT HYDROGEN

HELLIUM

FIRE (DISCRETE)

VE

INTEGRITY SWITCH REMAINS CLOSED

R

H

EA T

(A VE R

AG

E)

ALARM SWITCH CLOSES

O

SENSOR PRESSURE (PSIA)

40

NONHAZARDOUS OPERATING CONDITION

20

-65

250

390

1,000

EXPOSURE TEMPERATURE (° F)

Figure 26-5. Fire Detector Operating Principle (Functional)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Responder

NOTES

A hermetically sealed responder unit (Figure 26-3) is mated to one end of each sensor loop by a ceramic isolator that provides for passage of gas into the responder. The responder contains alarm pressure and integrity pressure switches (Figure 26-4) and a 3,000ohm resistor connected to and in series with the integrity switch. Ducts in the responder direct gas from the loop to one side of both pressure switches. The switches contain metal diaphragms that snap overcenter at different pressures. Under pressure, the diaphragm makes contact with a stationary pin to complete a circuit (switch closed). The switches are wired to a connector at the other end of the responder, mating with circuitry from the fire detection control unit. The integrity switch in the responder detects loop damage. It is normally held closed by gas pressure in the loop. From the switch, a 3,000-ohm resistor creates a voltage drop and passes the integrity signal to the control unit. If loop pressure is lost due to a leak or a cut in the loop, the switch opens, breaking the integrity signal and triggering the detection system fault indicators. The alarm switch signals the control unit if a fire or overheat condition occurs in the nacelle. The normally open switch snaps closed when sensor loop gas pressure increases beyond a point corresponding to an average temperature of approximately 390° F (199° C) or a local temperature of 1,000° F (538° C).

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ELECT EQUIP PANEL NO. 2

3010-M1 NO. 1 ENGINE FIRE DETECTION CONTROL UNIT

NO. 2 EQUIPMENT PANEL

2611-CU1 NO. 2 ENGINE FIRE DETECTION CONTROL UNIT

9811-J104

2611-CU2

Figure 26-6. Fire Detection Control Units

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Fire Detection Control Unit

NOTES

Each fire detection control unit (Figure 26-6) consists of a solid-state monitor circuit for each loop and a test circuit. Each circuit monitors its respective loop for both alarm and fault conditions. The test circuit provides self-test to the control unit by simulating both conditions in the monitor circuits. The loop monitor circuits are identical in logic operation. Fire alarm condition is monitored by a PULL FUEL OFF handle drive circuit and an alarm generator circuit. The handle drive circuit, upon receiving an alarm signal, applies power to illuminate the PULL FUEL OFF handle on the fire protection panel and the CHECK FIRE DET light. Upon receiving an alarm signal, the alarm generator circuit triggers the fire warning drive circuit, which in turn illuminates both ENGINE FIRE warning lights and sounds the fire bell (SOO 8105). Fault conditions developing from the integrity switch in the responder are monitored by the fault comparator and drive. Upon receiving a fault signal from the responder, the fault comparator and drive illuminates the respective loop FAULT light and the CHECK FIRE DET light. The test circuit contains an oscillator, which, when triggered by an external test switch, feeds the fault and test drive circuits. These circuits alternately trigger alarm test loops A and B alarm test circuits, providing the test signals for their respective PULL FUEL OFF handle drives and alarm generators. The circuits also alternately test the power failure and integrity faults for both loops through the fault comparator and drive circuit for each loop. The control unit also contains a fire warning cutout/reset circuit. When triggered by an external voltage from the ENGINE FIRE PRESS TO RESET switch, this circuit blocks the output of the triggered fire warning drive, extinguishes the ENGINE FIRE warning light(s), and mutes the fire bell (SOO 8105).

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FUEL VALVE OPEN CLOSED

EXTG AFT BTL

FUEL VALVE OPEN CLOSED

FUEL CLOSED

EXTG AFT BTL

VALVE OPEN

BAGGAGE SMOKE WARNING

BTL

FWD BTL

FAULT

FIRE

FIRE TEST

EXTG

TEST 1

FWD BTL BOTH

A

BOTH A

B

FAULT A

FAULT B

FAULT A

LOOP SELECTION

OFF FAULT B

TEST 2

LOOP SELECTION

ENGINE 1 PULL FUEL OFF

B

ENGINE 1 TEST DETECTION

PULL FUEL OFF

FIRE PROTECTION PANEL CHECK FIRE DET

ENGINE FIRE PRESS TO RESET

ENGINE FIRE PRESS TO RESET

WARNING PRESS TO RESET

Figure 26-7. Fire Protection Panel and Warning Lights

Figure 26-8. Glareshield and Master Warning Lights

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Fire Protection Panel

NOTES

The fire protection panel (Figure 26-7) features two LOOP SELECTOR switches, FAULT A and FAULT B lights, PULL FUEL OFF lights, and a TEST SELECTION switch. The LOOP SELECTION switches, one for each engine, have positions labeled “ A,’’ “BOTH,’’ and “B.’’ Switch positions A and B monitor loops A and B, respectively. With the switch in the BOTH position, both loops are monitored simultaneously. Four loop indicator lights provide visual indication of a fault in the loops or the associated circuitry. The FAULT A and FAULT B lights on the left side of the panel monitor the left nacelle; identical lights on the right side of the panel monitor the right nacelle. The PULL FUEL OFF handles, used for arming the fire-extinguishing system, illuminate to indicate a fire or overheat condition. The left handle serves the left engine; the right handle serves the right engine. Each handle has two bulbs installed behind a red lens and is powered by 28 VDC from the associated fire detection control unit. Electrical power is supplied to the LOOP SELECTION switches from the 28-VDC left and right essential buses through 5-amp circuit breakers. In the event one bus fails, the fire detection system for one loop in each nacelle continues to be powered by the other bus. Refer to Figure 26-9 to determine the power source for a particular loop.

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ALARM GENERATOR

FIRE WARNING DRIVE

A b B

MASTER FIRE WARNING

D

FIRE HANDLE FAULT A OUTPUT FAULT B OUTPUT CAUTION LIGHT

H F M G

}

28V DC (A) POWER IN FAULT COMPARATOR AND DRIVE

LOOP "B" ALARM

ALARM GENERATOR

PULL FUEL OFF HANDLE DRIVE LOOP B

POWER GND

R

DETECTOR B ALARM/INTEGRITY SIGNAL

c

B LOOP POWER (ONLY)

J

B LOOP POWER (BOTH)

K

A LOOP POWER (BOTH)

L

A LOOP POWER (ONLY)

C

TEST OUT RESET

E P

FIRE WARNING DRIVE 28V DC (B) POWER IN

FAULT COMPARATOR AND DRIVE

}

TEST CIRCUIT

MASTER FIRE WARNING CUTOUT/RESET

ALARM TEST LOOP B FAULT AND TEST DRIVE CIRCUITS ALARM TEST LOOP A

OSCILLATOR

TEST IN

N

}

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

PULL FUEL OFF HANDLE DRIVE LOOP A

DETECTOR A POWER DETECTOR A ALARM/INTEGRITY SIGNAL DETECTOR B POWER

FOR CONTINUATION SEE SHEET 2

26-12 LOOP "A" ALARM

ENGINE 1 FIRE DETECTION CONTROL UNIT

Canada Ltd.

Revision 2

Figure 26-9. Fire Detection System Electrical Schematic (Sheet 1 of 2)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

System Operation

NOTES

General The fire detection system monitors each nacelle, providing a visual indication of fire, overheat, or sensor loop fault. Since the system operation for each nacelle is identical, only the operation of the left nacelle system is described (Figures 26-9 and 26-10). Figure 26-8 shows the glareshield and master warning lights.

Fire Alarm Operation In the event of fire or overheat in the nacelle, sensor loops react to temperature increase with an increase in gas pressure. When the gas pressure reaches a preset value, the responder alarm switch snaps overcenter, closing the alarm switch. This causes a voltage increase of the detector alarm signal input to the control unit, triggering the PULL FUEL OFF handle drive and the alarm generator. The handle drive powers the PULL FUEL OFF handle lights, causing the handle to glow red and the CHECK FIRE DET warning light to illuminate. The alarm generator circuit triggers the fire warning drive, illuminating two flashing ENGINE FIRE warning lights on the glareshield panel and sounding the fire bell (SOO 8105). Pressing the ENGINE FIRE light applies a voltage to the fire warning cutout/reset circuit in the control unit, latching off the warning circuit and disabling the output to the ENGINE FIRE light(s) and silencing the fire bell (SOO 8105). The lights and the fire bell will then be available if a fire/overheat condition occurs in the other nacelle. The fire warning drive circuit can be latched off only if an alarm condition exists. If the alarm signal passes, the triggered fire warning drive circuit automatically resets for a subsequent alarm condition.

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300 SERIES SERIES 100 100 SERIES SERIES 300 ENG NO. NO. 11 ENG ENG ENG NO. NO. 11 ENG ENG NO. NO. 11 RESPONDER RESPONDER RESPONDER RESPONDER RESPONDER RESPONDER

LOOPA LOOP LOOP A

LOOP LOOP B B (PRE-MOD (PRE-MOD 8/1835) 8/1835)

LOOP LOOP B B AIR PIPE PIPE P2.5/3.0 P2.5/3.0 AIR (MOD (MOD 8/1835) 8/1835)

3K 3K A A D D

A A D D

B B C C

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL Canada Ltd.

Revision 2

Figure 26-9. Fire Detection System Electrical Schematic (Sheet 2 of 2)

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Loop Fault Operation If a sensor loop or its associated circuitry malfunctions, the FAULT lights, as applicable, illuminate simultaneously with the CHECK FIRE DET. Normally, current from the control unit is transmitted to the integrity switch in the responder. With the integrity switch in the normally closed position, current is routed through a resistor and back to the control unit through the detector alarm/integrity line to the fault comparator and drive. If either the detector power or the detector alarm signal wiring is broken, or if loop damage causes the integrity switch to open, the resistance of the responder increases, resulting in a voltage drop to the fault comparator and drive. If a short occurs between either line of the detector and ground, voltage to the fault comparator and drive decreases. When the voltage drops, the fault comparator and drive illuminates the CHECK FIRE DET light and the respective sensor loop FAULT light. A power failure to either loop control circuit is sensed by the respective fault comparator and drive. Alternate power to the failed loop fault comparator drive circuit is automatically provided by the bus of the functional loop. When the power failure fault is activated, the CHECK FIRE DET light illuminates simultaneously with the failed loop FAULT light. Power failure fault warning is available only when the LOOP SELECTOR switch is in the BOTH position.

TEST DETECTION switch is held. If only one loop is selected, its alarm circuits are tested continuously while integrity fault and power fault are tested alternately. During the test, the fire bell (SOO 8105) sounds and the following lights illuminate: • ENGINE FIRE warning light(s) flashing • CHECK FIRE DET warning light • PULL FUEL OFF handle light • FAULT A and FAULT B lights If the CHECK FIRE DET, PULL FUEL OFF, FAULT A, or FAULT B light does not illuminate or flash at 1/2-second intervals during the test, the associated loop has failed. Selecting the serviceable loop causes its FAULT light to illuminate steadily. Selecting the unserviceable loop causes its FAULT light to flash or not to illuminate.

NOTES

Fire Detection Test The test circuitry in the control unit becomes active when the TEST DETECTION switch is moved left or right to test a nacelle system. When initiating the test, power is applied to the oscillator circuit in the control unit. The oscillator feeds the fault drive circuits, causing them to alternately test the loop A and B fire alarms and integrity fault and power bus fault for both loops. The test signals simulate fault and alarm conditions in each loop circuit. The test continues to alternate as long as the

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

A

A

CLAMP ELECTRICAL WIRING HARNESS

RESPONDER

AIR SWITCHING VALVE

DIFFUSER PIPE

RH ENGINE VIEW

Figure 26-10. Fire Detection (300 Series) (Mod 8/1835)

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FIRE DETECTION— (300 SERIES)

NOTES

The fire detection system (Figure 26-10) on the 300 series aircraft works exactly the same as on the 100 series. On Mod 8/1835 aircraft, the existing loop B is disconnected from its control unit and capped off. A new loop B and responder is now connected to the control unit. The new loop B is mounted onto the right hand intercompressor case P2.5 pipe next to the air switching valve. References: • de Haviland, Service Bulletin 8-26-14 • Airworthiness Directive CF-91-26R1 • Pratt and Whitney Service Bulletin 21113

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ENGINE FIRE PRESS TO RESET

WARNING PRESS TO RESET

ENGINE FIRE PRESS TO RESET

FUEL EMERGENCY SHUTOFF VALVE POSITION INDICATOR LIGHTS ( GREEN ) (WHITE )

FIRE-EXTINGUISHING SELECTOR SWITCH

A

EXTG AFT BLT

FUEL OPEN

G A FIRE-EXTINGUISHING ARMING INDICATOR LIGHTS LOOP FAULT INDICATOR LIGHTS

VALVE CLOSE

W

FWD BLT BOTH A B

FAULT A

FIRE DETECTION LOOP ARMING SELECTOR SWITCH

FAULT B

LOOP SELECTION ENGINE 1 PULL FUEL OFF

TEST SWITCH

EMERGENCY FIRE CONTROL HANDLE AND LIGHT

Figure 26-11. Fire Protection Panel

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

ENGINE FIRE EXTINGUISHING

Fire Extinguisher Control Switches

General

Two EXTG switches, with positions labeled “AFT BTL’’ and “FWD BTL,’’ control the application of electrical power for detonation of the discharge squibs.

The engine fire-extinguishing system consists of two extinguisher bottles, discharge cartridges (squibs), and associated discharge valves and distribution lines to the nacelles. A red thermal discharge indicator provides visual indication of thermal discharge of the bottles; normal discharge of the bottles is indicated by a yellow indicator. A pressure switch/gage on each bottle completes a circuit to each arm advisory light to provide a visual check of bottle contents. The control circuit for each engine is armed by pulling the corresponding PULL FUEL OFF handle on the fire protection panel, which also initiates fuel and hydraulic fluid shutoff for the engine. Bottle discharge is controlled by EXTG switches on the panel. These switches initiate detonation of a corresponding squib on the appropriate bottle. Electrical power to the extinguishing system is from the battery bus. On airplanes with Mod 8/0195, additional circuit breakers provide an independent power supply to close the fuel emergency shutoff valves.

Fire-Extinguisher-Armed Indicator Lights Two amber extinguisher-armed indicator lights are located adjacent to each extinguisher switch. When the related PULL FUEL OFF handle is pulled, power is applied to the indicator circuits through the pressure switch/gage. The lights illuminate to indicate that the control circuits are armed and the associated bottle is charged.

Fuel Emergency Shutoff Valve Position Indicator Lights White FUEL VALVE CLOSED and green FUEL VALVE OPEN lights provide visual indication of fuel emergency shutoff valve position. Test circuits for the lights are powered from the advisory lights system.

NOTES

Components Description and Operation Fire Protection Panel The fire protection panel (Figure 26-11) incorporates the following switches and indicators.

PULL FUEL OFF Handles Pulling this handle completes circuits to arm the extinguisher control circuits and close the fuel and hydraulic fluid emergency shutoff valves. The handle illuminates when a fire or overheat condition is detected.

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

THERMAL RELIEF VALVE

DISCHARGE LINE TO NO. 2 NACELLE

DISCHARGE LINE TO NO. 1 NACELLE FILL AND THERMAL RELIEF VALVE

DISCHARGE VALVES AND SWIVELS ELECTRICAL CONNECTOR RECEPTACLE

THERMAL DISCHARGE INDICATOR (RED)

DISCHARGE VALVES AND SWIVELS

PRESSURE SWITCH /GAGE

SYSTEM DISCHARGE INDICATOR (YELLOW) PRESSURE SWITCH / GAGE (TYPICAL)

Figure 26-12. Fire Extinguisher Bottles Installation

FROM FIRE EXTINGUISHER BOTTLES

ZONE 3A

ZONE 2

DISCHARGE PORT (TYPICAL)

ZONE 3

ZONE 1

Figure 26-13. Distribution Lines

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

Fire Extinguisher Bottles The two stainless steel fire extinguisher bottles (Figure 26-12) incorporate discharge valves, a combination fill and thermal relief valve, and a pressure switch/gage. Each discharge valve consists of a plug that is held by a destructible housing and a stem. A swivel, capable of 360° rotation to facilitate tubing alignment, is located between the housing and the valve boss. The swivel incorporates an integral, threaded discharge port. An electrically actuated squib is installed in the housing. One discharge port on each bottle is connected to a discharge line to each nacelle. The fill and thermal relief valve consists of a fill fitting with a thermal relief valve connected in series. Under normal conditions, the valve functions as a check valve to prevent reverse flow through the fill fitting. If compartment temperature exceeds 187 ±10° F (86 ±5° C), bottle pressure rises to 1,300 to 1,450 psi, and a disc in the valve ruptures to relieve pressure through the fill fitting. Tubing connected to the fill fitting directs discharge to the red thermal discharge indicator. The pressure switch/gage provides bottle pressure reading. Inspection windows in the wing/fuselage fairing enable pressure checking without opening a panel. A pressure switch in the gage is connected to the bottle-armed light circuits. Contacts in the switch open as pressure decreases to 265 to 200 psi.

Each bottle has a volume of 378 cubic inches and is filled with 9 pounds of bromotrifluoromethane (CBrF3) (Halon). The bottles are pressurized with nitrogen to 600 to 625 psi at 70° F (21° C).

Distribution Lines The distribution lines from the discharge valve swivels on the extinguisher bottles terminate in designated fire zones (Figure 26-13). Tubing at discharge points is sized to deliver the correct quantity of extinguishant to suit the volume and ventilation characteristics of each zone.

System Discharge Indicator The yellow system discharge indicator (Figure 26-14), located on the fuselage skin, below the left inboard flap, provides visual indication of normal discharging of the bottles. The piston-operated indicator is connected through check valves and tubing to a tee in each discharge line. The discharge of either bottle into either discharge line acts on a piston to eject the yellow discharge indicator disc.

Thermal Discharge Indicator The red thermal discharge indicator, located above the system discharge indicator, provides visual indication of thermal discharge of the bottles. The indicator is connected through tubing to each bottle fill fitting. Discharge of either bottle due to high temperature in the compartment ejects the indicator and dumps the extinguishant overboard.

RED DURING VISUAL INSPECTION, IF THE DISK IS MISSING, THERMAL DISCHARGE HAS OCCURED. YELLOW DURING VISUAL INSPECTION, IF THE DISK IS MISSING, DISCHARGE INTO THE NACELLE HAS OCCURED.

Figure 26-14. Fire Extinguisher Discharge Indicator Discs Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

NO. 1 NACELLE

DISCHARGE SQUIBS (4) FORWARD EXTINGUISHED BOTTLE

TO RIGHT NACELLE

LEFT EXTINGUISHER DISCHARGE SWITCH

NO. 1 ENGINE FUEL VALVE POSITION INDICATOR LIGHTS

THREE-POSITION SWITCH IS SPRINGLOADED AT CENTER. AFTER PULLING PULL FUEL OFF HANDLE, MOVING SWITCH TO FWB BTL AND/OR AFT BTL DETONATES ASSOCIATED LEFT SQUIB, DISCHARGING ENTIRE CONTENTS OF SELECTED BOTTLE INTO NO. 1 NACELLE.

EXTNG AFT BTL

ILLUMINATE ALTERNATELY TO INDICATE FUEL VALVE IS OPEN (GREEN) OR CLOSED (WHITE)

AFT EXTINGUISHER BOTTLE

FUEL VALVE OPEN CLOSED

FUEL CLOSED

EXTC AFT BTL

VALVE OPEN

BAGGAGE SMOKE WARNING FWD BTL

FWD BTL BOTH

A

B

FAULT A

FAULT B

B

FAULT A

LOOP SELECTION

OFF FAULT B

TEST 2

LOOP SELECTION

ENGINE 1 PULL FUEL OFF

TEST 1

BOTH A

ENGINE 1 TEST DETECTION

EXTINGUISHER ARM LIGHTS (YELLOW) WHEN PULL FUEL OFF HANDLE IS PULLED. LIGHTS ILLUMINATE TO INDICATE ASSOCIATED AFT BOTTLE AND FORWARD BOTTLE ARE CHARGED AND ARMED FOR DISCHARGE INTO THE NO. 1 NACELLE. EACH LIGHT GOES OUT WHEN ITS BOTTLE IS DISCHARGED OR IF THE SYSTEM IS DISARMED.

PULL FUEL OFF

LEFT FUEL PULL OFF HANDLE WHEN PULLED, FUEL AND HYDRAULIC SHUTOFF VALVES TO NO. 1 ENGINE CLOSE, AND LEFT NACELLE DISCHARGE SQUIBS ARE ARMED FOR USE. PUSHING BACK IN DISARMS SQUIBS AND REOPENS SHUTOFF VALVES.

Figure 26-15. Engine Fire-Extinguishing System

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

System Operation

System operation with the PULL FUEL OFF handle pulled is shown in MSM Chapter 26.

General Engine fire extinguishing is manually initiated. The requirement for operation of the system for individual engine(s) is normally preceded by a warning from the engine fire detection system; however, a fire warning is not mandatory for operation of the system. With the master WARNING (red) light flashing, the CHECK FIRE DET warning light illuminates to indicate fire in one of the nacelles, and an illuminated PULL FUEL OFF handle indicates the affected nacelle. Fire extinguishing is initiated as follows.

NOTE Since operation of the fire extinguisher system is similar for both engines, only the operation of the No. 1 engine system is described.

Operating Sequence Pulling the PULL FUEL OFF handle (Figure 26-15) completes circuitry to initiate the following: 1. The No. 1 engine fuel emergency shutoff valve is energized closed, as indicated by illumination of a white FUEL VALVE CLOSED light. The green FUEL VALVE OPEN light goes out.

Selecting the EXTG switch to FWD BTL completes a circuit from the right essential bus to detonate the forward squib on the forward bottle, rupturing the destructible area of the discharge valve. The bottle discharges into the No. 1 nacelle, and the FWD BTL indicator light goes out as the bottle pressure decreases below 265 psi. Some of the discharged agent is routed through tubing to act on a piston that ruptures the yellow discharge indicator. Selecting the EXTG switch to AFT BTL, if required, completes a circuit from the right essential bus to detonate the aft squib on the aft bottle, rupturing the discharge valve. The bottle discharges into the No. 1 nacelle, as indicated by the AFT BTL indicator light going out. The armed (or actuated) extinguishing system is reset when the PULL FUEL OFF handle is pushed in (MSM Chapter 26), disarming the squibs and opening the fuel emergency shutoff valve and firewall hydraulic shutoff valve. The white FUEL VALVE CLOSED light goes out, and the green FUEL VALVE OPEN light illuminates.

2. The No. 1 engine hydraulic shutoff valve is energized closed. 3. The extinguisher switch is armed. 4. The forward squib of the forward fire bottle is armed and the No. 1 engine FWD BTL indicator light illuminates if the proper pressure is in the bottle. Power is from the battery bus. 5. The aft squib of the aft extinguisher bottle is armed and the No. 1 engine AFT BTL light illuminates if the proper pressure is in the bottle.

Revision 4—July 1995

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SMOKE DETECTOR INDICATOR LIGHT (RED) INTERRUPT SWITCH POWER INDICATOR LIGHT

1990 INTERIOR AND SUBSEQUENT

SELF TEST SWITCH

ALARM INDICATOR LIGHT

PINK RED

GREEN

GREEN

AMBER BLUE

SMOKE DETECTOR

LEGEND PRE-1990 INTERIOR

RED–LAVATORY SMOKE GREEN – (FLASHING) – OXYGEN PRESSURIZED DROP DOWN OXYGEN ONLY PINK – EMERGENCY CALL GREEN – P.A. ACTIVE AMBER – LAVATORY CALL BLUE – MASTER CALL WARDROBE

S1 A LAV SMOKE

NC C

DET OFF SW

G

C

NO NC

P2

NO 1 7 5

R

9 TO SERVICE CALL BUTTON IN LAVATORY

11

LAVATORY

Figure 26-16. Lavatory Smoke Detection System and Electrical Schematic

TECH C RR3321B

Canada Ltd.

Revision 4—July 1995

LAV 5 SMK DET 28-VDC R ESS BUS

TO PA SYSTEM HIGH CHIME

LAVATORY SMOKE DETECTOR

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

FLIGHT ATTENDANT'S ANNUNCIATOR PANEL

Canada Ltd.

DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

LAVATORY FIRE PROTECTION SYSTEM LAVATORY SMOKE DETECTION—MOD 8/0266 General The smoke detection system continuously monitors the lavatory to detect smoke generated by an incipient or active fire and provides visual and audible warning when smoke density exceeds a predetermined level. The system consists of a smoke detector mounted in the lavatory at the top of the forward wa l l , a wa r n i n g h o r n , a n d a t e s t s w i t c h (Figure 26-16). The detector is wired to a red indicator light on the overhead bin adjacent to the airstair door and activates an audible chime on the public address system. The detector contains a sensing chamber that houses an invisible infrared LED, a photodiode, and associated circuitry.

or until the LAV SMOKE DET OFF switch is selected to the off (amber) position. The audible sound from the detector ceases, the red indicator light goes out, and the chime on the public address system ceases.

System Test Selecting the LAV SMOKE DET OFF switch to the on (green) position and rotating the keyed test switch on the smoke detector 90° counterclockwise produces a simulated alarm condition to verify system operation. The audible warning sound from the smoke detector occurs simultaneously with the flashing red indicator light and the audible warning on the public address system. Rotation of the switch clockwise to its original position extinguishes the red indicator light and silences the warning from the detector and the public address system. Selecting the LAV SMOKE DET OFF switch off changes the switch coloring from green to amber.

NOTES

A push-button, alternate-action smoke detector switch above the wardrobe is labeled “LAV SMOKE DET OFF,’’ with amber lighting for the off position and green lighting to denote the on position.

System Operation Air in the lavatory passes through the smoke detector, entering the sensing chamber where it is sampled for the presence of smoke. Light from the pulsing LED light source is reflected by smoke particles into the photodiode light sensor. At the first detection of smoke, the pulse rating of the LED increases eight times above the normal rate. When the sensor confirms smoke for two consecutive pulses, it produces a signal to trigger an audible warning from the detector, illuminate the red indicator light, and produce an audible sound on the public address system. The smoke detector maintains the alarm signal until smoke density returns to a safe level

Revision 2

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DASH 8 SERIES 100/300 MAINTENANCE TRAINING MANUAL

INTERRUPT SWITCH POWER INDICATOR LIGHT

ALARM INDICATOR LIGHT

FRONT FLIGHT ATTENDANT'S ANNUNCIATOR PANEL

SELF-TEST SWITCH

SMOKE DETECTOR

D

FW

REAR FLIGHT ATTENDANT'S ANNUNCIATOR PANEL

D

FW

LAVATORY

AFT

Figure 26-17. Lavatory Smoke Detection System—Series 300

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LAVATORY SMOKE DETECTION—SERIES 300

NOTES

General The smoke detector system continuously monitors the lavatory compartment to detect smoke generated by an incipient fire and provides visual and audible warnings when smoke exceeds a predetermined level. The system includes one smoke detector assembly with a built-in evacuation horn, mounted in the lavatory at the top of the forward wall (Figure 26-17). The detector is interconnected with two red warning lights. One is located on the front flight attendant’s annunciator panel on the overhead bin beside the airstair door entrance. The other light is on the rear flight attendant’s annunciator panel in the galley area of the rear cabin. The smoke detector, mounted on a bracket, contains a dual-chamber ionization chamber sensing unit and associated circuitry.

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LAVATORY SMOKE DETECTOR

P3

8

3

6

12

14

7 R DS3 REAR FLIGHT ATTENDANT'S ANNUNCIATOR PANEL

R TO PA SYSTEM HIGH CHIME

R

DS2 FRONT FLIGHT ATTENDANT'S ANNUNCIATOR PANEL LAV SMK DET

R

5

28–VDC R ESS BUS

Figure 26-18. Lavatory Smoke Detection Electrical Schematic—Series 300

BOTTLE

NOZZLES

Figure 26-19. Lavatory Fire Extinguisher

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System Operation Air in the lavatory passes through the smoke detector and enters the dual-chamber ionization sensor. If smoke is present, the change in ion density is sensed, and a signal is generated that triggers an audible sound from the detector and illuminates a red alarm indicator light on the face of the detector. The two red warning lights on the flight attendant’s annunciator panels illuminate, and the public address system chime sounds once. The smoke detector maintains the audible alarm signal until the smoke density returns to a safe level or the interrupt switch on the face of the detector is depressed. The audible detector alarm and the chime are silenced, and the annunciator lights are extinguished. However, the red alarm indicator light on the face of the detector remains illuminated until the air clears, at which time it goes out, indicating the system has reset. An electrical schematic of the smoke detection system is shown in Figure 26-18.

The system is fully automatic. When a fire condition occurs and the temperature increases to approximately 170° F (76° C), the fused ends of the distribution lines melt, releasing the extinguishant into the garbage container.

NOTES

System Test Pushing the self-test switch on the face of the detector causes a simulated alarm condition which verifies operation of the system. The detector alarm indicator light and front and rear flight attendant’s annunciator lights illuminate, and the audible alarm horn and the public address system chime are activated. Releasing the self-test button or depressing the interrupt switch extinguishes the detector alarm indicator light and the annunciator panel warning lights. The detector alarm horn is also deactivated.

LAVATORY FIRE EXTINGUISHING The lavatory fire-extinguishing system consists of a fire-extinguishing bottle and associated distribution lines. The stainless steel, fully automatic, disposable extinguisher bottle with heat-activated, fused distribution lines is located in the lavatory garbage container (Figure 26-19).

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NO. 1 DETECTOR (CEILING MOUNTED)

NO. 2 DETECTOR (BULKHEAD MOUNTED) X543.50

DETECTOR NO. 2 (BULKHEAD MOUNTED)

X680.00

X610.00

X577.43

DETECTOR NO. 1 (CEILING MOUNTED)

SERIES 300

SERIES 100

EXTG AFT BTL

VALVE OPEN

A

BAGGAGE SMOKE WARNING TEST 1

FWD BTL BOTH B

OFF

FAULT B

TEST 2

SMOKE DETECTOR TEST SWITCH

LOOP SELECTION ENGINE 1

SMOKE

PULL FUEL OFF

TECH CHECK RR3322B 26-20

Figure 26-20. Baggage Compartment Smoke Detection System—Series 100 and 300 NO. 2 DETECTOR

X577.43

NO. 1 DETECTOR X680.00

BULKHEAD-MOUNTED SMOKE DETECTOR P4

A B C D EF

P3

A B C D E F

1J3

A B C D EF

BAGGAGE COMPT CEILING-MOUNTED SMOKE DETECTOR P2

A B C DE F

4

5 6

P1 1 2 3 4 5 6 7 8 9 10 11 12

5 SMOKE DET

SMOKE DETECTOR CONTROL AMPLIFIER SIGNAL IMPUT

1 2

3

S1 TEST SW

TEST 2

TEST 1

BEACON ALARM GROUND 28 VDC TESTING

OFF

LOGIC CIRCUIT

SMOKE MASTER CONTROL PANEL

28 VDC R ESS BUS PRE MOD 8/0235

DIM AND TEST LOGIC

SMOKE

28 V DC

WARNING LIGHTS PANEL

TECH CHECK RR3321B 26-21

Figure 26-21. Baggage Compartment Smoke Detection System Electrical Schematic

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BAGGAGE COMPARTMENT SMOKE DETECTION SYSTEM— SERIES 100/300 GENERAL The smoke detection system continuously monitors the rear baggage compartment to detect smoke generated by an incipient or active fire, providing a visual warning when smoke density exceeds a predetermined level. The system includes two identical smoke detectors, one on the bulkhead and one on the ceiling, a control amplifier on the avionics rack, and a test switch (Figure 26-20). Each detector contains a labyrinth that houses a beacon light to provide a constant light source, a photocell, a test light, and associated circuitry. The components are shock-mounted to protect against vibration, and a perforated, removable cage protects them from damage. The detectors transmit electrical signals to the amplifier, which illuminates the red SMOKE warning light as required.

nate the SMOKE warning light and flash the master WARNING light. The detectors maintain the alarm signal until density returns to a safe level, at which time the voltage to the amplifier decreases to 9 to 11 VDC and the amplifier resets and extinguishes the warning lights.

SYSTEM TEST With the test switch held in the TEST 1 position, the No. 1 smoke detector test light comes on to provide a light source directed at the photocell. This causes a simulated alarm condition to verify operation of the system. The SMOKE warning light comes on, and the master WARNING light flashes. Releasing the switch to OFF extinguishes the warning lights. Holding the switch at TEST 2 repeats the above results for the No. 2 smoke detector. Figure 26-21 is an electrical schematic of the baggage compartment smoke detection system.

The system is powered by 28 VDC from the right essential bus through the 5-amp SMOKE DET circuit breaker.

SYSTEM OPERATION Air in the baggage compartment passes through the perforated cage of the smoke detectors and enters the labyrinth where it is sampled for the presence of smoke. Smoke in the airflow directs light from the beacon light into the photocells, lowering their resistance. This initiates a voltage proportional to smoke density into the control amplifier. If the smoke is sufficient to reduce light transmission in the labyrinth to 90 ±5%, the voltage from one or more of the detectors exceeds the amplifier alarm point of 13 to 15 VDC. This causes an alarm relay in the amplifier to energize, triggering an electronic logic circuit to illumi-

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BAGGAGE COMPARTMENT DIVIDER BULKHEAD (LOOKING AFT) INDICATOR DISC

GREEN

RED YELLOW

AFT OF PILOT'S SEAT FLIGHT ATTENDANT'S CLOSET Figure 26-22. Portable Fire Extinguishers—Series 100

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PORTABLE FIRE EXTINGUISHERS— SERIES 100

NOTES

Three manually operated fire extinguishers are installed in the airplane, one aft of the pilot’s seat and two in the cabin (Figure 26-22). The extinguishers in the cabin are located in the emergency storage compartment and on the divider bulkhead in the baggage compartment. All of the extinguishers are retained in quickrelease brackets, and instructions for their use are displayed on the extinguisher. The Halon 1211 universal extinguishant in each bottle is suitable for use on electrical, fuel, or oil fires. It is not corrosive and does not cause cold burns, harm fabrics, or leave residue.

PORTABLE FIRE EXTINGUISHERS— SERIES 300 Four manually operated portable fire extinguishers are provided: one aft of the pilot’s seat, one in the emergency stowage compartment, one on the baggage compartment divider bulkhead, and one behind the front row rear facing passenger seat on the right side.

NOTE Portable fire extinguishers may be replaced by different types by a customer, but locations will remain the same.

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MAINTENANCE CONSIDERATIONS MAINTENANCE PRACTICES The maintenance practices found in the Maintenance Manual usually relate to removal/installation and adjustment/test. The following procedures apply to the fire protection system in a general sense: • The clearance between the fire detection sensor and the adjacent structure must not be less than 0.25 inch. • Exercise extreme care during maintenance not to twist, kink, or dent the sensor loop. Bend radii must be constant and no less than one inch.

• If a sensor loop connector is contaminated, clean with a solvent approved for use with electrical connectors, and blow out with dry air or nitrogen. Never use any lubricant in connectors. Never use carbon tetrachloride as a cleaner. • Use a light source and a mirror to check that the blanking or shunting wire is removed from cartridges prior to installing electrical connectors. • Prior to assembly and tightening fire extinguisher system tubing, apply a light c o a t i n g o f F E L P RO - C 6 2 1 t o m a l e threads at each connection, with the exception of the first two threads entering the female threads.

WARNING

• Fire-extinguishing agent (bromotrifluoromethane) is nonpoisonous but semitoxic. Contact with the agent or breathing its vapors should be avoided. Thoroughly ventilate closed areas before entering.

Fire extinguisher bottle squibs are pyrotechnic devices. Inadvertent detonation can cause injury to servicing personnel. The electrical connector pins on the squib must be protected by a blank or shorted by a piece of wire.

• Never attempt to repair a sensor loop or a detection control unit in the field. They are remove-and-replace items if they fail or are damaged.

WARNING

• If a sensor loop is stored, ensure that it is kept in an area where no physical damage will occur to the outer shell. Do not place objects on top of a sensor loop. • The number of times a sensor loop is bent for installation must be kept to a minimum to avoid cold working of the outer shell. • When working close to a sensor loop, or on the unit itself, take precaution to prevent damage to the loop by tools and equipment (scratching the outer shell, denting, crushing, etc.).

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Do not loosen cartridge housing assembly when removing or installing a cartridge. The bottle may discharge and cause serious injury to servicing personnel.

WARNING Do not use a voltmeter, flashlight battery, continuity light, or any similar device to test fire bottle cartridges. These devices may affect the useful life of the cartridge and could cause inadvertent detonation.

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SERVICING The nacelle fire extinguisher bottles cannot be recharged in the airplane. Removal and replacement with a serviceable bottle is necessary since bottle recharging requires special equipment available only at an authorized agency equipped to perform recharging. The bottles are charged with bromotrifluoromethane and are pressurized with nitrogen. The portable fire extinguishers are charged with Halon 1211 agent and are nitrogen-pressurized; the extinguishers can be recharged at any approved fire equipment service shop.

INSPECTIONS Unscheduled Maintenance Checks Chapter/Section 5-50-00 of the Maintenance Manual provides complete instructions for maintenance checks (inspections) required after overstress or abnormal use of specified systems. It does not contain any inspections pertaining to the fire protection system. However, Chapter/Section 26-20-11 provides an inspection/check of the fire extinguisher bottles.

FUNCTIONAL CHECKS Functional (operational) checks are normally performed after maintenance has been accomplished on a system. They may also be performed when a system is suspected of malfunctioning. Functional checks are frequently a part of a maintenance procedure and are not called out separately. They sometimes include such maintenance tasks as final adjustment, torquing, and safetying. Chapter 26 of the Maintenance Manual provides procedures for performing the following functional checks: CHAPTER/SECTION • Engine fire detection system ........................................26-1100 • Smoke detection system

26-12-00

• Engine fire-extinguishing circuits .......................................26-2000 Due to length and complexity, these checks are not presented in this chapter. Refer to the Maintenance Manual chapter/section listed.

Scheduled Maintenance Checks Scheduled maintenance checks are listed in the Maintenance Program, PSM 1-8-7.

FAULT ANALYSIS Isolation of a fault or malfunction can be accomplished by a systematic analysis of the trouble, beginning with the most probable cause and progressing to the least probable cause. Any system(s) interfaced with the troubled system should be operating properly prior to troubleshooting.

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