Avionics - Systems And Troubleshooting Cap. 6

Chapter 6

Section 1

Learning Objectives:

Introduction

• Autopilot Theory

Autopilots have been installed on aircraft for severa! decades. The systems ha ve been proven reliable and, for the most part, more accurate than human pilots. The variety of autopilot systems is almost as vast as the variety of airplanes. Light aircraft may have simple autopilots installed; while transport category aircraft often incorporate complex systems with full autoland capabilities. Most twin-engine aircraft incorporate sorne type of autopilot system and many corporate aircraft employ complex systems similar to large passenger jets. Autopilot technologies w ill also play a large role in the upcoming decades as airspace becomes more crowded dueto the increased number of flights globally. The initiative known as NEXTGEN, or Next Generation Air Transportation system, relies on modern technologies to enhance safety and improve capacity. Autopilots were first developed to relieve the pilot/co-pilot from constantly having to ha ndle the aircraft controls. On long flights this beca me especially important on older transport category aircraft like the Boeing 707. These aircraft were difficult to control and were physically exhausting, especially in bad weather. As technology improved, and systems became lighter and smaller, autopilots began to filter into the light aircraft market. In the early 1980s, autopilot technology advanced to the point where the machine became more efficient than the human; flying with advanced autopilot system s saved both time and fue!. Automated flight systems developed in the 1990s and early 21st century, improved safety while increasing the number of aircraft in the authorized airspace. Today a

•Autopilot Components •Air Data • Compass Systems •lnertia/ Reference Systems •lnertial Navigation Systems

Left. Modern autoflight systems are a far cry from the original autopilots of the 1930s. Today's systems, with featu res like autoland, can almost operate the aircraft fro m the departure gate to the arrival gate.

6-2

PILOTVIEWS VARIOUS INSTRUMENTS TO DETERMINE --AIRCRAFT ANO NAVIGATIONAL INPUTS

Autopilot and Autojlight Systems

t 1

lot, and presents severa[ modern systems in depth. Many modern autopilots are known as autoflight systems, both terms can be considered synonymous; however, it is generally accepted that autopilots are relatively simple and autofligh t systems are more complex and capable of more func tions.

HUMAN PILOT

---+1

VARIOUS AIRCRAFT 1------. CONTROLS

~

1 1

1 1 1 1

PILOT MOVES CONTROL WHEEL. RUDDER PEOALS, ANO THROTTLE AS NEEOEO

L---------- ~E~O_!l~C_!S -- - ---- -- J PILOT SENSES AIRCRAFT MOVEMENT ANO AOJUSTS CONTROLS ACCOROINGLY

AUTO PILOT

Sectíon 2 CONTROL SURFACES

SERVOS

r---· ~

L-----~

Basic Autopilot Theory

1

1

wide range of autopilots is available for almost any type of aircraft.

By definition, the autopilot is designed to perform pilot duties automatically. The autopilot must first interpret the aircraft's current attitude, speed, and location. Second, if adjustments are necessary, the autopilot must move the appropriate control su rface, and throttles on advanced systems. Third, the autopilot m ust anticípate the aircraft's movement and reposition the control surfaces, and/or throttles, to prevent the aircraft from overshooting the desired course and attitude.

Simple autopilots provide guidance along only the longitudinal axis of the aircraft. Thesc systems, often found on light single-engine aircraft, were called wing levelers because they were used to keep the wings level. More complex systems provide total control for attitude and navigation. Many modern transport category aircraft incorp orate systems that provide aircraft control, yaw damping, navigational guidance, thrust control, and Category III landing capability. This chapter examines the operational theory of autopilot systems, explains individual subsystems of the autopi-

To p erform the functions just described, the autopilot mu st monitor various aircraft parameters including airspeed, altitude, pitch, roll, and yaw (Figure 6-2-1). Navigational aides are also monitored to provide course data. ext, the autopilot wil~ analyze the data to decide if adjustments are needed. An autopilot computer is used to analyze the data and output the necessar y control infor mation. If adjustment is required, ser vos are u sed to move control surfaces and reposition the aircraft. Servas are devices used to move the flight controls, or throttles, in accordance with autopilot com-

1 1 1 1 1

1

~-------------------COMPUTER SENSES AIRCRAFT--MOVEMENT ANO~ AOJUSTS CONTROLS ACCORDINGLY

Figure 6-2-1. Comparison of human pilot and autopilot

Yaw Channel

AIRCRAFT INPUTS · Air Data .

~Vertical Speed Airspeed

-1 1

1

_L

Pitch Channel

Altitude Heading Navigation data Attitude (gyro, or INS)

1

1

- ¡1

Roll Channel

1 Pilot's 1 Copilots Autopilots Disconnect Switches

•

1

L---- --- -- -- --- ----- -------~ FOLLOWUP

Flight Director Command Bars HSI Display (Optional)

Figure 6-2-2. Basic autopilot block diagram

Autopilot and Autoflight Systems mand signals. Hence, older autopilot computers were sometimes referred to as servo amplifiers. A follow-up system is used to inform the autopilot computer that the control surface has changed position. The follow-up system allows the computer to anticípate when the control surfaces should be returned to the neutral position. Since aircraft move in three axes (pitch, roll, and yaw), many autopilot systems typically contain three channels or subsystems. The autopilot must be capable of performing all of the necessary functions in a safe and reliable way. Two major safety considerations that ever y autopilot system must have, is the ability to be quickly and positively disengaged electronically, and the ability to be overridden manually by the pilot, if necessary, to regain control of the aircraft.

1

6-3

(A)

o

HDG o NAV

(B)

Figure 6-2-3. Typical autopilot control panels from a corporate-type aircraft: The autopilot block diagram (Figure 6-2-2) shows the inputs and outputs for a typical three-axis autopilot. The autopilot computer receives aircraft inputs from: l. Air data sources supplied by pitot static

pressures, or electronic signals from an air data computer 2. Heading sources provided by the aircraft's compass system 3. Navigational inputs, such as ILS, DME, or VOR

4. Attitude information from an inertial reference system (IRS) or attitude sensors A quick disconnect push-b utton is on the control yoke to force a quick disconnect of the autopilot. Each autopilot system w ill incorporate sorne means by which the flight crew can input commands. Thc two panels show n in Figure 6-2-3A are used to control an autopilot for a typical corporate jet aircraft. The top panel is u sed to engage the autopilot and to control the manual pitch and roll functions. This panel also conta.i ns the AP XFR, for transfer of a dual autopilot system, and turbulence mode (TURB) push-button. The panel shown in Figure 6-2-3B is used to select the different modes of the autopilot system. From this panel the pilot can select severa! different modes of operation including navigation (NAV) or vertical navigation (VNAV). The navigation mode allows the autopilot to fly the selected lateral navigational course (i.e., control of north, south, east, and west direction s). The vertical navígation mode allows the autopilot to fly the selected altitude or glide path. The ma in autopilot outputs include the three sign als used to control the pitch, roll, and yaw servos. As the respective control surfaces

(A) Used to engage the autopilot and fo r manual pitch and rol! functions, (B) Used to select the different modes of the autopilot system Courtesy of Rockwe/1 lnternational, Collins Divisions

move according to autopilot signals, a followup signa! is transmitted back to the autopi lot computer. The autopilot computer may also have an output dedicated to the flight director control. The flight director presents a visu al aid to the pilot that is used for manual control of the aircraft. As seen in Figure 6-2-4, the flight director display is typically incorporated in the attitude director indicator (ADI) or integrated into the primary flight display (PFD). Since the flight director utilizes many of the same inputs as the autopilot, the two systems often share components. In the exam ple of Figure 6-2-2, the autopilot computer drives the command bars for the flight director display.

FLIGHT DIRECTOR COMMAND BARS

Figure 6-2-4. A typical flight di rector display incorporated on an attitude director indicator (ADI)

6-4

1

Autopilot and Autoflight Systems

Yaw Damping Virtually all high-speed aircraft are desig ned w ith swept back wings. The aerodynamics of a swept back wing causes a s tabi lity problem known as Dutch rol/. Dutch rol! is basically a slow o scillation of the a ircraft about its ve rtical axis. Correct application of the rudder can prevent dutch roll; however, it requires constant repositioning of the r udd er. This process becomes almost impossible for the pilot. The yaw damper system is d esigned to control rudder and eliminate dutch rol!. The yaw damper is basically an autopilot component d edicated to rudder control. If the system d etects a slip or skid of the aircraft, the rudder is activated to correct this condition; hence, eliminating dutch rol!. On most a ircraft, the yaw damper is considered independent of the autopilot system, although they may share the same control p an-

z

els, sensors, and compu ters. Du ring flight, the autopilot and yaw damper may both be used to position the rudder. The autopilot positions the rudder to coordinate turn activity; the yaw damper positions the rudder to eliminate dutch roll. On most aircraft, the yaw damper can be engaged independent of the autopilot.

Section 3

Autopilot Components A variety of components are incorporated into every autopilot system. Many of these components or subsystems are not actually part of the autopilot, but are essential to autopilot operation. Older au topilots employ analog systems and m echan ical sensors. Newer au topilot components are typically digital systems and communicate th rough data bus cables. The following presen ts m any of the basic elements of an autopi lot system.

Gy roscopic Sensors There are two common t ypes of g yroscopic sensor s u sed in modern auto pilots: rotating mass gyros and ring laser gyros. Most autopilots use gyro sensors to detect movement of t he aircraft. Gyro ou t puts are also used for re ference on certain navigation systems. Gyro systems a re both fragi le and expensive; it is very important the technician becomes familiar w ith the system before perform ing maintenance on gyros.

Rotating Mass Gyros

OUTPUT SIGNAL PROPORTIONAL TO GYRO MOVEMENT

Rotating mass gyros have a tend ency to stay stable in space. This effect allows a rotating m ass gyro equipped w ith a rate sensor to detect aircraft motion (Figure 6-3-1). The output signa] from the rate sensor can be sent to an autopilot computer and/or used to stabilize a gimbal platform .

Fig ure 6 -3-1 . Diagram of a basic rotating mass gyro

STABLE PLATFORM (ACCELERATION SENSORS

MOUNTED HERE) ~

3 GYROS (ONE FOR EACH AIRCRAFT AXIS)

Figure 6-3-2. Diagram of a typical gimbal platform

Gimbal platforms. Rotating mass gyros are often used to stabilize acceleration sensors m ounted on g imbal platforms. A gimbal platform is made rig id in space, parallel to the earth's surface regardless of the aircraft's atti-tude. This is accomplished b y mounting three rotating gyros to the platform as show n ir: Figu re 6-3-2. O ne gyro is needed for each axis of the gimbal. If the aircraft's attitude changes. the gyr o r ate sensor pr oduces an output signa: that is propor tional to the amount of attitude change. This sig na ) is monitored and ampli-

Autopilot and Autoflight Sysl"ems

1

6-5

fied by the autopilot computer, then sent to the torquer units on the stable platform. The torquer produces the counter-force, typically a magnetic field, needed to stabilize the platform. With the acceleration sensor mounted on a stable platform, an accurate acceleration can be monitored by the system. If the acceleration sensors were not mounted to a gimbal platform, the sensors would measure attit ude changes as well as aircraft accelerations. Due to advancements in modern technologies, rotating mass gyros and gimbal platforms are quickly being replaced by more reliable and accurate systems.

Laser Gyros The r ing laser gyro (RLG) is actually an angular rate sensor and nota gyro in the true sense of the word. Conventional gyros generate a gyroscopic stability th rough the use of a spinning mass. The gyro's stability is then u sed to detect aircraft motion. The RLG uses changes in light frequency to measu re angular displacement. The term laser stands fo r "light amplification by stimulated emission of radiation." The RLG system utilizes a helium-neon laser; that is, the laser's light beam is produced through the ionization of a helium-neon gas combination. A typical RLG is shown in Figure 6-3-3. This system produces two laser beams from the same source and circulates them in a contra-rotating triangular path. As shown in Figure 6-3-4, the high voltage potential, approximately 3,500 V, between the anodes and cathodes produce two light beams traveling in opposite directions. The laser is housed in a glass case, wh ich is drilled with precise holes to allow travel of the light beam. Mirrors are used to reflect each beam around an enclosed triangu lar area. A prism and detector are installed in one corner of the triangle. The prism reflects the light to allow both laser beams to be measured by the detector. The resonant frequency of a contained laser is a function of its optical path length. When the RLG is at rest, the two beams have equal travel d.istances and identical frequencies. When the RLG is subjected to an angular d isplacement around an axis and perpendicular to the plane of the two beams, one beam has a greater optical path and the other has a shorter optical path. Therefore, the two resonant frequenáes of the individuallaser beam change. This change in frequency is measured by photosen;;ors and converted to a d igital signal. Since the frequency change is proportional to the ang ular displacement of the unit, the system's dig ital output signa! is a direct function of the angular rate of rotation of the RLG.

Figure 6-3-3. A ring laser gyro

READOUT DETECTOR

CORNER PRISM

Figure 6-3-4. Pictorial diagram of a ring laser gyro

The RLG system is typically coupled to a complete navigation system. The digital signals from the RLG can be used to control inertial reference and navigation systems and/or autopilot functions. Each inertial sensor assembly contains three triangular lasers. In Figure 6-3-5, two of the lasers can be seen mounted w ithin the sensor assembly. RGL technology provides a ncw era of aircraft safety. Since the RLG has no moving parts, it has a much greater reliability than the conventional rotating mass gyro system. No matter which type of gyro is em ployed, all three axes of tbe aircraft must be monitored for a fully functional autopilot. Pitch, roll, and yaw can only be monitored by sensors aligned in the correct position. On most systems, alignment

Co urtesy of Honeywell, ln c.

6-6

Autopilot and Autoflight Systems

1

units are interchangeable between locations on the aircraft. Pin programming is used to identify the specific installation location .

Figure 6-3-5. An inertial sensor assembly containing three ring laser gyros Courtesy of Honeywe/1, lnc.

FIXED POINT ON AIRCRAFT

<:l===J FORWARD

o

-2 -1

+1 +2

ACCELERATION SCALE

BASIC PENDULUM (Al

FORWARD

~

~Q 2 2 Q QQ QQ

CENTERI NG SPRING (2 EA.)

J.--.--,----.-.--.--, 1

: _

2

:

:

_

1

:

O +

+ 1 2

HOUSING FIXED TO AIRCRAFT STRUCTURE

ACCELERATION SCALE

BASIC ACCELEROMETER (B)

Figure 6-3-6. Simple pendulum used to measure acceleration: (A) A basic pendulu m, (B) A basic aircraft accelerometer becomes very important during the instaJiation, removal, and replacement of rate sensing LRUs. Each unit must be oriented in the correct position with respect to the rest of the aircraft. For example, if a unit is installed 180° out of alignment, the system could react exactly opposite of what is necessary. Be su re the unit is installed according to the manufacturer's instructions. Many gyro

The most modern ring laser gyros op erate on the same principies as those found in 10-year -old aircraft. The main improvements in modern RLGs is the miniaturization of components allowing for the construction of lighter and smaller gyros, which in turn has allowed manufactu rers to integrate gyr o components in a single housing. For example, air data and inertial reference units are now commonly combined into an air data initial reference unit (ADIRU). Another advancement found in modern laser gyro systems is the ability to electronically align the unit. In these systems, when installing the gyro, all final align ment is performed through electronic circuitry during initial set-up operations. The technician must simply follow the required steps and the RLG softwa re will ensure that the system is aligned with the aircraft and the vertical, longitudinal, and horizontal axes.

Maintenance and Troubleshooting In general, laser gyro systems are relatively maintenance free and rotating mass gyros seem to be subject to frequent failures. Every gyro provides sorne type of electrical output signa!. The simplest way to detect the operation of a gyro is to check the output signa!. On the most modern systems, the output is probably a digital signal. On many rotating mass gyros, the ou tput is a three-phase AC sign a!. Many rotating mass gyros are combined with rate sensors. The output signa! from the rate sensor is typically a single AC voltage. On any gyro system, the output signals are relatively low current; therefore, a poor connection can easily create a loss of output. If a gyro's output is inaccurate or missing, check the electrical connection s to the unit. Rotating mass gyros al! produce a humming noise as the gyro rotates. This noise should be present any time the gyro is active. If the gyro is completely quiet, the unit is defective or there is no power to the gyro assembly. If the gyro assembly makes an unusually loud rumble noise, the gyro bearings are probably worn beyond limits and the unit should be replaced. All gyros require sorne warm-up period. Laser gyros must reach a given temperature to stabilize; rotating mass gyros must reach a certain RPM, and digital control circuits often perform test functions prior to operation. In many gyro systems, it is important that the aircraft remain stationary during this warm-up or initialization p eriod. Be sure to provide sufficient warmup time to all gyro systems whenever performing maintenance.

Autopilot and Autoflight Systems

l\cceleronneters An accelerometer is a device that senses ai rcraft acceleration. Acceleration is a vector force and therefore is measured in both magnitude and direction. Since an aircraft can accelerate in three directions, a mínimum of three accelerometers are used on most installations. Figure 6-3-6A depicts how a basic accelerometer resembles a simple pendulum. The pendulum will swing to the right as the aircraft moves forward and to the left as the airCl·aft decelerates. Figure 6-3-6B shows the arrangement of a simple aircraft accelerometer. This design incorporates two springs to center the pendulum. As the aircraft moves, the indicator moves in the opposite direction, relative to the aircraft. An act ual accelerometer incorporates a pick-off device to convert pendulum movement into an electric signa! (Figure 6-3-7). The electric signa! is amplified and, in sorne cases, converted to a digital format. Sorne accelerometers u se a secondary current flow to keep the armature centered over the pick-off coils. A torquer current is used to produce a magnetic field that centers the armature in the null position. This produces a n armature that is very stable and increases accelerometer accuracy. ~odern accelerometers are often small microdectromechanical systems (MEMS) consisting of little more than a cantilever beam with a small mass at the end. This mass is often called the !JToof mass. Under the influence of externa! accelerations the proof mass deflects from its neutral position. An electronic element is used :o sense the deflection and produce an analog or digital signa!. This method is simple, reliable, and inexpensive. Micromechanical accelerometers are available in a wide variety of measuring ranges. The designer must make a .:ompromise between sen sitivity and the maximum acceleration force that can be measured.

TO ARMATURE TORQUER

---1

·

NULL SIGNAL IF USED

1

6-7

1-- ----. ·

SPRING HOUSING FIXED TOAIRCRAFT STRUCTURE

FORWARD

PENDULUM

'--------

-

EXCITATION INPUT CURRENT

Figure 6-3-7. An ai rcraft accelerometer showing the pick-off un it

Figure 6-3-8. A MEMS accelerometer

1ost micromechanical accelerometers operate m-plane, that is, they are designed to be sensidw to accelerations in only one plane. By inte5Jating two devices perpendicularly on a sinde die, a two-axis acce!erometer can be made. 3y adding an additional out-of-plane device, 2rree axes can be measured. Such a combina;;:on always has a much lower misalignment ::rror than three discrete models constructed :x:lependently and mounted into an assembly.

microns (20 millionths of a meter) to as large as a mill imeter. Figure 6-3-8 shows components of a typical MEMS accelerometer. MEMS accelerometers are found in a variety of applications on aircraft and other consumer products. For example, automobile airbag controls use a MEMS device to measure sudden decelerations, smart phones use MEMS to determine if the display is horizontal or vertical and laptop computers use MEMS accelerometers to instantly park the hard drive if a fall is detected.

fEMS accelerometers are designed in a manner smilar to microelectron ic circuits, such as, micro-:roc:essors. MEMS devices are constructed of silicon using a photo-etching process, which has =roven to produce high quality components with !:eat reliability at low cost. A complete MEMS D:elerometer measures typically in a range of 20

Modern accelerometers typically contain one of two electronic elemen ts: piezoelectric or capacitive components, which conver t mechanical motion into an electrical signa!. Piezoelectric accelerometers rely on crystals to produce a voltage/current when pressure is applied to the cr ystal. The vol tage/current

EN SOR

To PITOT SYSTEM

ÜUTLET

TOTALAIR TEMPERATURE PROBE

PJTOTTUBE

Figure 6-4-1 . Typicallocations for installation of pitot tubes, static ports, and temperature probes

produced is directly proportional to the fo rce applied to the crystal, within the limits. If the crystal is made small enough, an acceleration force "pressing" on the crystal will create an electrical output. The electrical output is directly proportional to the pressure created by the acceleration. This is an ·extremely weak analog signa! that is typically amplified and

AIRSPEED INDICATOR

ALTIMETER

PITOT PRESS URE

STATIC PRESSURE

Capaciti\·e accelerometers typically use a silicon micromachined sensing element. Like a!: capacitors, the capacitance va lue will change according to the distance between the plates.. The sensor is constructed of an extremer small capacitor and, like the crystal acceleiometer; an acceleration force easily affects i:.. As the aircraft accelerates the microcapacito: plates bend changing the distance bet\\·ee;:; the plates. The plates on this capacitor woulC be measu red in m icron s and thus invisib.e to the naked eye. The m in iatu re size m ake: the capacitor vulnerable to even the slightes< acceleration force. The accelerometer circuit:rr measures this change in capacitance, w hich iS directly proportional to the acceleration ~ creates the appropriate output sig na!. On m any transport category aircraft, there are at least two accelerometers to monitor the acceleration of each aircraft axis. The outputs of both accelerometers are combined to provide an extremely accurate measurement The accelerometers for each axis may also be mounted at opposite ends of the air craft (i.e.. the tail and the nose). This also improv~ accuracy. Acceler ometers are used in a variety of appücations on modern aircraft. Originally accelerometers were fou nd in autopilo t circuits used for nav igation and to determine aiicraft position. Eventually as aircraft systerns became more integrated and accelerometers became smaller, lighter, and increased sensitivity, engineers found additional uses for these sensors. For example, today modern aircraft employ accelerometers to measure sudden position cha nges caused by turbulence. A n accelerometer can de tect a sudden l ift allowi ng a computer to actívate flight controls, such as spoilers, and change wing efficiency lift. This creates a smoother ride for passengers and decreases stress on the aircraft stru cture.

Sectíon 4 VERTICAL SPEED INDICATOR TEMPERATURE PROBE

ÜUTSIDE AIR TEMPERATURE

Figure 6-4-2. A pneumatic pitot-static air data system

Air Data To fly safely, it is essential to know the aircraft's airspeed and the current altitude. These items are difficult for the pilot to sense and are best-determined using flight instrum ents. To measure these basic flight parameters, the air mass surrounding the aircraft must be monitored. The measurement of this air mass is known as air data. The

Autopilot and Autojlíght Systems three air data elements ty pically measured are temperature, static pressure, and pitot pressure. There are two different temperatures typically measured by an air data system: static air temperature (SAT) and true air temperature (TAT). Static air temperature is the temperature of the undisturbed air surrounding the aircraft. True air temperature is a measure of the air as it is comp ressed by the moving aircraft. Temperatures are an important air data reference used to improve the accuracy of other parameters and enhance the efficiency of modern au toflight systems. Static p ressure. Static pressure is the absolute pressure of the air that surrounds the aircraft. Static pressure varíes inversely w ith the altitude of the aircraft and also changes with the general atmospheric conditions of the area. On a stand ard day, 59°F at sea leve!, t he static pressure is 29.92 Hg (1013 mb). Static air pressure should be measured in undisturbed air, which is difficult to find near a moving aircraft; therefore, correction factors that are calculated b y the a ir data computer are often employed when determ ining static pressure. Static pressure is used to determine the aircraft's altit ude and vertical speed.

Pitot pressure. Pitot pressure is an absolute pressure of the air that enter s the pitot tube. With the a ircraft at r est the pitot pressure is equal to static pressu re. Since the opening of the pitot tube faces the direction of aircraft travet as the aircraft increases speed pitot pressure will increase. The d ifference between pitot pressure and static pressure is often referred to

as dynamic pressure. Dy na mic pressu re is used to determine the aircraft's airspeed. Figure 6-4-1 shows the installation locations of typical static ports and pitot probes. Both pitot a nd temperatu re probes are typically heated to prevent ice formation. Redundant probes may be installed on the opposite side of the a ircraft to avoid errors caused by aircraft yaw.

Types of Air Data Systems There are basically three ty pes of air data systems curren tly in u se: pneumatic, electropneumatic, and electron_ic. Each of these systems can be connected to an autopiiot. In generat it might be said that newer air craft d esign s incorporate more electronic air data systems while older and less complex aircraft employ pneumatic systems. A pure pneumatic system relies solely on the static and pitot pressures to drive the alti m eter and vertical speed indicators (Figure 6-4-2). Pneumatic air data system s can be used only on simple autopilots. The electropneumatic air data system employs both electronic circuitry and pneu matic functions (Figure 6-4-3). The pitot and static air pressures are sent directly toan altimeter and vertical speed indicator s that are often used as backup instruments. Both pitot and static pressures are also sent to the electronic air data unit along with sign als from the air temperatu re sen sors. The electronic air data unit converts the pressure d ata into an electrical signa! and then sends that information to one or more displays and the auto pilot computer. On large

A IRSPEED INDICATOR

ALT ALERT CAUTION: ANNUNCIATOR

ALTI METER

VERTICAL SPEED INDICATOR

A IR D ATA

L.::=====~ ElECTRONICS TEMPERATURE--------------~~

UNIT

PRO BE

ÜUTSIDE A IR TEMPERATU RE

=igure 6-4 -3. Block diagram of a typical electropn eumatic air d ata system

1

6-9

6-1 O 1 Autopilot and Autoflight Systems

corporate or transport aircraft, electronic signals may also be transmitted to other aircraft systems, such as the flight data recorder orcentral maintenance computer. Figure 6-4-4 shows a simplified example of how a modern light aircraft might employ an electropneumatic air data system.

PRIMARY FUGHT DISPLAY (PFD)

-SHOWING AIRSPEED, ALTITUDE, VERTICAL SPEED, ANO 0THER AIR DATA

PITOT PRESSURE ......,._ __

Electronic air data systems are employed on the newest aircraft and typically convert all air pressure values to an electrical signa! for distribution. There are no pneumatically operated instruments on this type of system; even the backup altimeter, airspeed, and vertical speed indicators are electronic instruments (Figure 6-4-5). This design concept greatly reduces weight, especially on large aircraft, since all plumbing lines used to distribute pitot and static pressures are kept to a mínimum. Electrical w iring or data bus cable is used to distribute the air data information to all end users. The electronic unit employed in this type of system is typically called an air data computer (ADC).

___. AIR DATA ElECTRONICS UNIT

Fig ure 6-4-4. A modern air data system interfaced with an electronic flatpanel display

To AUTOPILOT PITOT PRES SU RE

C1!!1~==::::l

TO ÜTHER A IRCRAFT SYSTEMS

Ali-80A BAROMETRIC ALnMETER

ASI· 800 AIR SPEED INOICATOR

TO WARNING ANNUNCIATOR

AIRSPEED INDICATOR PRE·BOA PRESELECTORIALERTER MSI·SOF MACH SPEED INOICATOR

STATIC

PRE SS U RE C1iil~=::::l

AIR DATA (OMPUTER

AlTI METER TAI·SOA TEMPfTAS INDICATOR

' \ ' 1

., 12 /> 4 ;_· CLINt

· ::- =

6-

, 1 2 4,/

TEMPERATURE PROS E

V ERTICAL SPEED INDICATOR

1 \

TEMPERATURE INDICATORS

Figure 6-4-6. Components for a typical corporate aircraft air data system Figure 6-4-5. Distribution of air data signals to the various aircraft systems

Courtesy of Rockwelllnternationa/, Collins Divisiam

Autopilot and Autoflight Systems

BB EJ BB AIR DATA TO AUTOPILOT ANO ÜTHER SVSTEMS

-

-1

- PITOT PROBE 1 - STATIC PORT 1 - AIR TEMPERATURE SENSOR$

AIR DATA TO r---•AUTOPILOT ANO ÜTHER SYSTE MS

- PITOT PROBE 2 - STATIC PORT 2 - AIR TEMPERATURE SENSORS

Figu re 6-4-7. An air data system interface with a typical integrated display system

Air Data Computer Systems An ADC system monitors pitot pressure, static pressure, and air temperature to determine various parameters. The air temperature input is most likely an analog electric signal produced by a temperature transducer. For most systems, the ADC wou Id receive pitot, static, and air temperature inputs from redundant sources. The ADC will transform the pressure inputs to an electrical signa! u sing pressure transducers. The electrical signals are then sent to the processing circuitry where the data is manipulated into useful information, such as indicated airspeed, vertical speed, altitude, etc. The ADC outputs whether either analog or digital depending on the system are sent to the autopilot, flight instruments, flight warning systems, central maintenance computer, and other systems that require air data information. The components of a typical air data system employed in corporate-type aircraft are shown in Figure 6-4-6. This system would be found in an aircraft employing individual electromechanical instruments. Aircraft with integrated display systems would send digital air data signals to sorne form of display management computer where they are processed and sent to the pilot's and co-pilot's PFD (Figure 6-4-7). As discussed in Chapter 3, these types of systems are often highly integrated,

employ digital data transfer systems, and many of the components share software and /or operational functions.

Maintenance and Troubleshooting Modern air data systems incorporate built-in diagnostics that can be accessed through the aircraft's central maintenance computer system or through a specific LRU. If electromechanical instruments are used to display air data, they often employ a self-test button. Pressing the self-test button will cause a specific indication on the instrument. If this indication is not displayed, the instrument should be replaced. lf the indicator passes the self-test, then there is most likely something wrong w ith the electronic air data unit, related wiring, or the pitot/ static plumbing. On aircraft equipped w ith central maintenance systems, air data faults wou Id be stored and accessed through the central maintenance com puter. In many cases, built-in diagnostics can detect the fault on the first attem pt. However, even on modern systems there are many mechanical faults to the system plumbing that cannot be detected by the diagnostics. Faults with partially

1

6-11

6-12

1

Autopilot and Autoflight Systems

Figure 6-4-8. An air data test unit clogged pitot tubes or static ports or leaking plumbing lines can give inaccurate readings on air data instruments. Static ports are especially prone to clogs since their openings are relatively small. Bent or misaligned pitot tubes will also create accuracy problems. Water that may have entered the system can also cause problems. On dual systems, if only one probe is damaged the ADC may reporta disagree message on the fault diagnostics. Any time an air data system malfunctions, be sure to inspect the static and pitot probes for damage or clogs. Modern airspace around the world has become more crowded and regulatory agencies have reduced separation minimums between aircraft to help improve traffic control. In the United States, the FAA has a standard known as Reduced Vertical Separation Minimums (RVSM) that are requ ired for most high altitude flights. The accuracy of air data instruments is critica! in order to safely reduce vertical clearance. RSVM aircraft are required to perform air data accuracy checks according to the FARs every 12 calendar months. There are severa! commercial air data test units available from various manufacturers. A typical air data test unit is shown in Figure 6-4-8. In general, each tester will ensure accurate airspeed, vertical speed, altitude, and air temperature data. They w ill also test for pitot/ static leaks and integrity. Many test systems w ill ensure the interface reliability between air data equipment and various end users.

Section 5

Compass Systems Any autopilot that perfonn s basic navigation functions must receive data from the aircraft's compass system. The magnetic compass is too inaccurate; therefore, the autopilot system

Fig ure 6-5-1. A flux detector typical of those found on corporate aircraft

must rely on an electrica l/electronic compass system. The flu xgate Compass system employs one or more remote sensors to produce an electric signa! th at can be used to determine the aircraft's position relative to magnetic north. The newest aircraft integrated u nits, such as Honeywell Aerospace's Attitude Heading Reference System (AHRS), supply inertial reference data, magnetic heading information, and air data combined into one unit. Honeywell's AHRS employs MEMS technology accelerators, and integrates attitude sensors and digital laser gyros; eliminating the need for a fluxgate coropass and performing in-flight alignment verification using a GPS cross reference. An AHRS unit of this type weighs about 10 pounds and uses approximately 20 watts of electrical power. Another common integrated system is known as the ADIRU (air data iner tial reference unit). This system combines air data sensors and the components of a typical IRU to provide a variety of cr itica! fligh t data including airspeed and heading information. In most cases each of these integrated systems work in conjunction with a com plete flight management and/ or autoflight system. The remote sensor used in a fluxgate coropass system is often called a flux detector or flux valve (Figure 6-5-1). The flux detector receives a constant input of 115 YAC, or 28 YAC, at 400 Hz. The output voltage is a function of the al ignment of the detector w ith the earth's magnetic field. The sensing unit in the flux detector is the flux valve. The flux valve is comprised of a three-spoked frame w ith an output winding on each spoke (Fig ure 6-5-2). An excitation winding is located in the center of the flu x frame. The frame is typically suspended in a sealed case on a universal joint, wh ich allows it to pivot and rem ain relatively stable at different aircraft attitudes. The u nit

Autopilot and Autoflight Systems

1

6-13

EXCITATION OUTPUT WINDINGS

~CO I L

(400 HZ INPUT)

AIR DATA (OMPUTER 2

MAGNETOMETER

MOUNTING FLANGE (2 EA.) FLUX FRAME (b)

GIA INTEGRATED AVIONICS UNIT 2

Figure 6-5-2. Flux detector interna! components: (A) The flux trame, (B) The flux trame and housing assembly ATTITUDE HEADING REFERENCE SVSTEM 2

is surrounded by oil to dampen the flu x frame movement. The operation of the flux detector relies on the interaction of the earth's magnetic field and the magnetic field induced in the flux frame by the 400 Hz excitation coil. Without the earth's magnetic field, each output coil of the flu x frame would produce an equal voltage. As the aircraft moves with r espect to mag netic north, different legs of the flux frame becom e saturated with magnetism. As the saturation of the fram e changes, different voltages will be induced in the three output coils. Ther efore, the output coils produce a three-phase AC voltage that changes characteristics r elative to the aircraft's heading. In older systems the ou tput signals from the flux detector can be sent directly to a rem ate slaved compass such as those found in horizontal situation indicators (HSI). This requires a relatively large current flow in order to move the HSI; therefore the flu xgate itself must also be large. On more modern systems, the flux detector output sig nals may be sent to an electronic circuit where they are amplified and distributed to various systems th at will u se the data. This allows for the desig n of a smaller and more sensitive fluxgate. Systems that m onitor flux detector signals include the autopilot system, the flight data recorder, the radio m agnetic indicator (RMI), and / or the electronic flight instrument systems (EFIS). O n sorne aircra ft, the output of the flux detector is sent to a coropass coupler. The compass coupler contains a mechanical ser vo /synchro combination . The output from the compass coupler is directed through relays to the RMI or HSI. The relays are u sed for reversionary sw itching to select

~ HIGHSPEED DATA BUS

~ RS-485

~

~

ARINC429

Figure 6-5-3. The magnetometer sig nal is converted into di gital information on most modern systems.

which source would be used to drive which RMI or HSI. As w ith many modern aircraft systems the r em ote compass or magnetometer, is often combined with other functions. As seen in Figure 6-5-3, the magnetometer in the Garmin G-1000 system converts al! information to a d igital data format. The mag netic information is sent to the AHRS unit in an RS-485 format; AHRS converts the signa! to ARINC 429 data and send s compass signals to the PFD a nd the integrated avionics unit (IAU). The IAU uses compass data for a variety of navigation and autopilot functions.

RS-232

6-14

1

Autopilot and Autojlight Systems CAUTION: When testing any fluxgate compass system, be sure the aircraft is away from items that may interfere with the earth's magnetíc field . The test must be done outsíde the hangar and away from other aircraft, cars, railroad tracks, and power cables. Also be aware that many metal items are not visible, such as buríed power lines, fue/ tanks, or concrete reinforcing rods. Placing the aircraft on a compass rose is the best way to test the system. These general practices must also be observed by the pilot during a prejlight test. If a problem occurs with the fluxgate system on the ground, be sure to retest the compass system in a known environment.

A

A HC-85

FDU-70 FLUX DETECTOR

Figure 6-5 -4. Typical insta llation locations for a flux detector and attitude Courtesy of Rockwelllnternotional, Cal/ins Divisions head ing computer

Com pass system tes ts should be perfor med with various electrical equipment both on and off. If the slaved compass is affected by the operation of cer tain electr ical equ ipment the problem m ust be fixed. First check to see if al! wires near the flux detector are properly shielded. If proper shielding fails to produce the desired results, reroute electr ical wires or equipment away from the flux detector to illuminate the error.

Maintenance and Troubleshooting In most installation s, the flu x detector is placed in the aircraft's w ing to isolate it from magnetic interference cau sed by the other electrical systems (Figure 6-5-4). The mounting structure for the flux detector contains an adjustment that is used to en sure the unit is correctly aligned. This align ment becomes an important mainten ance item when replacing the unit. In general, alignment of the flux detector is done by placing the aircraft facing a known com pass point and moving the flux detector until the remote compass on the H SI or PFD reads the correct compass heading. Th is is a simplified explanation of the procedures since most systems contain two flu x detectors. Always follow the manufacturer's recommended procedures for flux detector alignment. To isolate a defective component in the compass system, test all LRUs containing built-in test equipment. Troubleshooting the RMI or HSI typically becomes a remove and replace procedure. That is, the suspect comp onent is swapped with a known operable u n it and the system is tested. The flux valve un it is a nonrepairable item and must be replaced if fo und to be defective. Keep in mind the flu x valve (magnetometer) creates a relatively weak electrical signa!. Dirty, worn, or loose electrical connector pins can easily affect weak signals. Be s ure all connectors are in good condition whenever troubleshooting compass systems.

Section 6 In ertial R ef erence Systems An inertial reference system (IRS) is a combination of laser gyros and accelerometers u sed to sense the aircraft's angular rates and accelerations. IRSs are relatively expensive and typically fo und only on corporate, transport, or military-type aircraft. The LRUs of a ty pical inertial reference system are shown in Figure 6-6-1. The laser gyros and accelerometers are installed in the inertial reference un it (IRU), w hich is typically installed in the aircraft's equipment bay. The IRU also contains the computer circui try for signa! processing and system interfacing. The data produced by an IRS is u sed in conjunction with a total au toflight system. The IRS data is typically combined with air data outputs to compute: l. Attitude (pitch, roll, yaw)

2. Angular rate changes (pitch, rol!, yaw) 3. Aircraft velocity 4. Course track angle 5. Inertial altitude 6. Linear accelerations

Autopilot and Autoflight Systems

1

6-15

7. Magnetic heading

8. True heading 9. Position (Jatitude and longitude) 10. Vertical speed 11. Wind speed 12. Wind direction The output data from the IRS is a primary input for a modern autoflight system. IRS outputs are also sent to electronic flight instrument systems for display of attitude and navigational data. IRS data is sent to the flight data recorder along w ith other aircraft systems. Many of the la test IRSs are so accurate, the need for a flu xgate compass is eliminated. For example, aircraft such as the B-757/767, the B-747-400, and the A-320 use the IRS for magnetic heading data. The IRS sends magnetic compass data to the RMI and /or EFIS for display to the flight crew. Sorne state-of-the-art inertial reference systems integrate IRS with magnetic heading and air data functions. The advanced Attitude Heading Reference System (AHRS), by Honeywell Aerospace, integrates attitude sensors and laser gyros. This system uses MEMS technology accelerators and advanced digital circuitry to miniaturize component size and save weight. The AHRS un it elimina tes the need for a fluxgate compass and performs in-flight alignment verification using a GPS cross reference. Th e Honeywell system weighs approximately 10 pounds and uses approximately 20 watts of electrical power. Another common integrated system is known as an air data inertial reference unit (ADIRU). This system combines a ir data sensors and the components of a typical IRU to provide a variety of critica! flight data including air speed and heading information. In most cases, integrated systems work in conjunction with a complete flight management and/or auto flight system.

Initialization Since an IRS can only measure changes in position, the unit must be given a starting reference ?Oint. The procedure used to provide the IRS ·ith initiallatitude and longitude is called ini'ialization. Initia lization typically occurs at the aircraft gate befare the first flight of the day. If :he aircraft has not been moved overnight, the ,?OSition in memory can be used. If the aircraft nas been towed to a new location, the crew :r~ust enter the correct latitude and longitude ;nto the IRS, typically using a multifunction .Iphanumeric keyboard. :>uring initialization, the IRS accelerometers the direction of the earth's gravity

=:~easure

Figure 6-6-1. Components of a typical laser inertial reference system Courtesy of Honeywe/1, lnc.

Figure 6-6-2. Local ve rtical is measured between the aircraft's location and the rotational axis of the earth. force to determine the aircraft's local vertical. Local vertical is a direction perpendicular to the rotational axis of the earth that intersects the aircraft's position (Figure 6-6-2). During initialization, the IRS rate sensors measure the speed and direction of the earth's rotation relative to the a ircraft. This, along with the latitude, longitude, and local vertical allows the system to determine true north. At the completion of the initialization process, the IRS computer contains the necessary data to compute the aircraft's current position and heading. Initialization ta kes approximately five to ten minutes and the aircraft cannot be moved during this time.

6-16

1

A utopilot and Autoflight Systems

lATERAl AXIS

"-

l ONG ITU DINAL

AXIS /

(A)

RO LL

,.......,

i: \\ !

\'..--

(B) YAW

(C)

PITCH

Figure 6-6-3. The three axis of the aircraft; one IRS unit must be aligned with each axis

Theory of Operation Each IRS unit is made up of three laser gyr os and three accelerometers. One each of these un its is aligned with the pitch, roll, and yaw ax is of the aircraft (Figu re 6-6-3). Figure 6-6-4 shows the three laser gy ro assemblies and accelerometers w ithin a typical IRU. The three gyros measure ang ular displacement about their resp ective axis (pitch, rol!, and yaw). The accelerometers are u sed to measure the rate of acceleration about each axis. Each of the three axes must be monitored since the aircraft travels in three-dimensional space. Also, most aircraft will contain two or three

JRUs, each wi th the capability to monitor aJI three axes of the aircra ft. Multiple IRUs provide the redundancy needed for safety anri reliability. Once the IRU has been initialized, the system knows wh ere it is located in all three dimensions and current head ing. As the aircraft moves in any d irection from its initial position, the IRS will sense the movemen: a nd compu te the new location and heading u sing a high s peed processor. Using Figure 6-6-5 as an example, ass ume the aircraft is s tationary at a given point in space (Point A If the longitudinal accelerometer measu res

Autopilot arzd Autojlight Systems

ACCELEROHETERS - -

ROLL RATE

PITCif RATE

SENSJNG

SENSING

LASER GYRO

LASER GYRO f'").

AIR OATA

X .....-:J

COHPUT ERS - ~~> ·

GYROS ·- ·-· -

COHPUTER

LONGITUOAL

AXIS
YAW RATE SENSING LASER GYRO

z AXIS VERTICAL (YAW)

FLIGIIT

HANAGEHENT COMPUTERS - ~~>·

INERTJAL REFERENCE UNIT

Figure 6-6-4. The typical configuration of an inertial reference unit.

STATIONARY AIRCRAFT

~INT

ú'~

A)

AIRCRAFT VELOCITY 20 FT/SEC

(P-+~+IN_T_B_)

ACC:::::T-IO_N_ _

Courtesy of Northwest Airlines, tnc.

AIRCRAFT VELOCITY 20 FT/SEC LOADED 2,000 FT PAST POINT B

_ _ _A_C_C_ E::::,:" O

(POI~T

C)

OF 2FT/ SEC 2 2,000 FT DIFFERENCE

Figure 6-6-5. The IRS measures acceleration and time to calculate the aircraft's change of position. a n acceleration of 2 ft /sec2; this mean the aircraft is accelerating for ward. After ten seconds the aircraft would be flying ata velocity of 20 ft /sec (10 x 2 ft /sec 2 = 20 ft /sec). Assume the aircraft stops accelerating and the velocity remains constant at Point B. If the aircraft continues to fly with a velocity of 20 ft/sec for 100 seconds the aircraft's new location (Point C) is 2,000 feet from Point B. Distance equals velocity multiplied by time (20 ft /s x 100 = 2,000 ft).

constantly being compared on multiple IRU systems to ensure accuracy.

The IRS computer performs similar calculations for the angular rate changes measured by the laser gyros. Assuming the IRS detects a yaw rate of 5° per second for 15 seconds, the computer would determine that the heading has changed 75° from the aircraft's original heading. The IRS computer continuously performs acceleration and angular rate calculations for aH three axes. By measuring both accelerations and angular rates, the IRS can provide a constant update on the aircraft's location and head ing. Heading and location information are

3. Laser gyro is subject to drift over time.

Severa! other factors that can affect the accuracy of the IRS are the: l. Earth's r.otation at approximately 15.04°

per hour 2. Spherical shape of the earth meaning air-

craft do not travel in a straight Jine over the surface This drift is much less than rotating mass gyros; however, it is still important to consider To compensate for these inherent errors, the TRS software is programmed to make the necessary corrections while processing the data. On many systems, another means of ensuring accuracy is to periodically cross-reference location data with other navigational aids, s uch as GPS.

1

6-17

6-18

1

Autopilot and Autoflight Systems

~

CONTRAL SURFACE CABLE

~

CONN ECTION OF BRIDLE TO MAIN CABLE

Figure 6-7-1. Diagram of a simple pneumatic servo

Maintenance and Troubleshooting Most aircraft that employ an IRS also contain sorne type of centralized maintenance computer system (CMCS). IRS troubleshooting is typically accomplished using this system. Faults are stored in a nonvolatile memory and displayed when requested by the technician. In most cases, the aircraft will contain two or three IRUs. Each of these units is interchangeable and can be swapped to help identify a defective unit, or reversionary switching can be used to swap units from the flight deck. Whenever removing any IRU, make sure to handle the unit with care; the IRUs are fragile and can be severely d amaged if dropped. Any unit that has been dropped is not airworthy. Whenever shipping an IRU~ be sure to use the appropriate shipping container to help protect the unit. Whenever testing an IRS, be sure to allow proper time for the aircraft to initialize. The aircraft must remain stable during the initialization process. Large w ind gu sts or maintenance being performed on the aircraft may upset the initialization procedure. In this case, the procedure should be repeated on a "quiet" aircraft. Never condemn an IRS that w ill not align on the first attempt, it may be caused by a moving aircraft. Installation of the IRS in the correct position w ith respect to aircraft axes is also an important consideration. For example, if the IRS unit was installed with the gyros and accelerometers out of alignment, the system could not produce accurate data. This is typically not a problem for removal and replacement since the LRU is installed in a rack that is permanently mounted to the aircraft. However, if the rack should become bent, cracked, or somehow misaligned, the IRS will not work properly. Sorne modern IRUs can be electronically aligned. In this case, the physical alignment is less critica! and the fine adjustment is made using software corrections.

Section 7

Inertial Navigation Systems A modern inertia l n avigation system (I S) uses airborne eq uipment for aircraft navigation without relying on externa! radio signals. Laser gyros and accelerometers provide three d imensional navigation capabilities. This system is em ployed mainly by the military. The major advantage of the system is that it requires no externa! navigation aid s. All the equipment for wor ld w ide navigation is contained in the INS. Many older tran sport category aircraft, such as the DC-10, early B-747s, and the L-1011 employ an INS which utilizes rotating gyros, gimbal platforms, and accelerometers to sense aircraft position. This system is called 11\5 since it is resp onsible for coordinating the aircraft's navigationa l parameters, including flight plan a nd wayp oint selection. The I -s can be interfaced w ith the autopilot or flight di rector system to steer the aircraft. The d ifference between a nd INS and IRS is that the INS provides waypoint and fl ight plan capabilities; a modern day IRS must work in conjunction with a flight management system (FMS) to provide these functions.

Servos A ser vo is a device used to apply a force to the aircraft's control surface in response to an auto-pilot command. There are basically three types of servos: pneumatic, electrical, and hydraulic. There are also hybrid servos; these are typicalt combinations using hydraulic activators combined with an electric motor. Each servo mt& incorporate sorne type of mechanism so the pilo;: - can override the autopilot command. They alsc typically contain a feedback system that pffr' vides a return signa) to the autopilot com puteL

•

Autopilot and Autoflight Systems

Pneumatic Servas

Hydraulic Servas

Pneumatic servos are vacuum actuated units used on simple autopilots for light aircraft. As seen in Figure 6-7-1, the pneumatic servo operates using a vacuum applied to the ser vo diaphragm. The autopilot computer controls the vacuum. Two servos are required for each control surface. A bridle cable is u sed to connect the servo to the control surface cable. Pneumatic ser vos offer limited range of travel and provide a relatively weak actuating force; therefore, pneumatic ser vos have limited use. Today pneumatic servos are only found on older aircraft u sing simple autopilot systems.

Hydraulic ser vos are the mos t powcrful type of ser vo actuator; hence, these units are typically used on tra nsport category aircraft. Ever since the B-727 and DC-9 took to the ai r, tra nsp ort category aircraft have employed hyd raulically operated control surfaces. In brief, the systems operate using an engine-driven hyd raulic pump and employ a control valvc to route hyd raulic fluid to a control surface actuator. The control surface actuator is mechanically l.in ked to the control surface. On most transport category airC1·aft, the control surface actuator is linked to the control wheel and r udder peda ls through control cables. On the newest aircraft, like the A-380 and B-787, the control whccl or side stick controller and rudder pedals are connected to the control surfaces via electrical wiring and computer circuits.

Electric Servas Electric ser vos utilize an electric motor and clutch assembly to move the aircraft's control surface according to autopilot commands (Figure 6-7-2). Due to their reliability and excellent torque production, electric servos are commonly found on all types of aircraft, including trainers, personnel, corporate-type turbine and turboprop aircraft. A bridle cable is installed between the servo's capstan assembly and the control surface cable. The capstan is used to w ind / unw ind the bridle cable; hence, moving the control surface. Figure 6-7-3 shows a ca pstan w ith the test fixtures installed in preparation for adjustment of the slip clutch assembly. This figure also shows the torque adjustment nut for the sli p clutch . The slip clutch gives the pilot the ability to overpower the unit in the event of a servo malfunction. Most electric servos use sorne type of clutch assembly to connect the servo motor to the capstan. During manual operation of the controls, the clutch is d isengaged and the capstan moves freely. Dur ing autopilot operation the clutch is engaged and the capstan is connected to the servo motor. The clutch engage/disengage is accomplished using an electromagnet simu lator to a solenoid. The clutch is engaged when the electromagnet is energized and the servo is active. The autopilot computers and pilot activated controls are used to energize the servo clutch . Virtually all autopilot systems employ an autopilot disconnect switch that is typically a push-button located on the control wheel. This gives the pilot the ability to instantly disconnect the autopilot function and manually fly the aircraft. The electric motor and clutch assembly usually operate on direct current; however, sorne transport category aircraft may employ AC motors. In order to keep capstan r.p.m. relatively low while the motor operates at high speed, the unit contains a gear reduction assembly. The gear assembly is normally self-lubricating and requires no regular maintenance.

1

6-19

Figure 6-7-2. A typical electric servo Courtesy of Rockwe/1 lnternational, Collins Divisions

The A irbus A-320 hydraulic servos op erate in two different modes: active and damping. As seen in Figure 6-7-4, the active mode is employed when the servo valve is pressurized and the Elevator Aileron Computer (ELAC) energizes the solenoid valve. In the active mode, hydraulic fluid is controlled by the servo valve and directed through the mode selector valve to the aileron act uator. In the damping mode, the actuator follows control surface movement as hydraulic fluid is allowed to flow through the restricting orífice. The servo is in damping mode whenever the solenoid is de-energized or hydraulic pressure is not supplied to the servo valve. In this

HOLDDOWN SCREW

•

CAP STAN ..--- HOLDING FIXTURE

CAPSTAN . / LOCKING ./" PEN CAPSTAN TEST FIXTURE

'"'~ WRENCH NUT "-._:

:....- .. .

TORQUE ADJUSTMENT NUT

Figure 6-7-3 . An electric servo capstan mounted to the ca pstan test fixture Cou rtesy of Rockwe/1/nternational, Collins Divisions

6-20

1

Autopilot and Autoflight Systems

(A)

KEY: HP=HIGH PRESSURE (SUPPLY) R=RETURN PRESSURE ELAC=ELEVATOR AILERON COMPUTER

which do not reguire a connection to the ai:"craft's main hydraulic system . An independer.! servo actuator is more reliable and allows the aircraft manufacturer to reduce or illuminatr the plumbing needed for a central hydraulic system. This reduces overall aircraft weig~ This type of servo receives an electronic signa! which controls an electric motor containea within the actuator assembly. The electric motor is directly coupled to a hydraulic pump which supplies the pressurized fluid; the actuator moves accordingly. This type of actuat
(B)

Figure 6-7-4. Diagram of a t ypical hydraulic servo and related control circuitry: (A) Active mode, (B) Damping mode

situation, the mode selector valve moves to the r ight via a spring force and connects the actuator hydraulic fluid to the restricting orífice. The mode selector valve and the main aileron actuator both produce a feedback signa! to the computer (ELAC 1). Although there are a large variety of hydraulic servos for transport category aircraft, they all operate in a similar fashion. Each hydraulic ser vo w ill contai n some type of servo or control valve, typically actuated by an electric solenoid. The control valve will move hydraulic fluid to the main actuator in order to move the control surface. Every unit w ill also contain sorne type of bypass or damping mode in the event of system failure. On fly-by-wire aircraft, th e flight deck controls are moved via an artificial force produced by the autoflight computer(s). On more traditional aircraft, the control wheel and rudder pedals move vía a cable linked to the actuators.

Hybrid Servas Some modern aircraft, such as the B-787 or A-380, employ servos that combine hydraulic actuators and electric drive motors. These electro-hydraulic servos are self contained units

Servo Feedback Systems As mentioned earlier, all ser vo systems must contai n sorne type of feedback circuit to inform the autopilot computer that the control surface has moved. The feedback system produces an electrical signa! that is directly proportional to the movement of the servo actuator. There are two common devices u sed to genera te the feedback signa]: an AC synchro and a differential transducer. Synchros are typically employed on electric and hydraulic ser vos used in conjunction with analog autopilot systems. State-ofthe-art digital autopilot systems often employ d ifferential transducer feedback systems.

Synchro Systems The most common autopilot feedback synchro is a tra nsformer-like device that monitors angular displacement using a stationary primary w inding and pivoting secondary winding. As shown in Figure 6-7-5, the primary winding receives a n input voltage of 26 VAC 400 Hz. The output voltage of the secondary is a function of the angular position of the secondary winding. Figure 6-7-6; position number one, shows the secondary w inding in the null position or perpendicular to the primary w inding. In this position, no voltage is induced in the secondary. As the secondary rotates clockwise, the voltage induced in the secondary increases until the secondary is parallel to the

Autopilot and Autoflight Systems primary (p osition number four). The voltage then decreases as the rotor continues to turn clockw ise. A second null is reached w hen the rotor becomes horizontal to the primary once again (position number seven). The secondary voltage is in phase w ith the p rimary voltage for rotor positions two through six.

1

6-21

PRIMARY WINDING (STATIONARY)

26V 400HZ AC INPUT ----+-

- - - - - L_

As shown in Figure 6-7-6, w hen the secondary w inding ro tates past the second null (p osition number 7), the output voltage is 180° out of phase with the primary voltage. The voltage value continues to change as the secondary continues to rotate clockwise. The out-of-phase condition exists until the rotor reaches the first null position once again . The output voltage is 180° out of phase w ith resp ect to the input voltage in positions eight th rough twelve.

_l

-f!m

PIVOT

¡---

MECHANICAL LINKAGE'TO SERVO ACTUATOROR CONTROL SURFACE

1

t

SECONDARY WIN DING (MOVABLE)

OUTPUT SIGNAL 400 HZ AC

Figure 6-7-5. Components of a typical autopilot feedback synchro

OUTPUT VOLTA GE IN PHASE WITH INPUT VOL TAGE

POSITIONS #1

#2

#3

#4

#5

#6

--,

_________________ ______ _ _______________________________ ___ ____________________ _____ J1

#7

#8

#9

# 10

#11

-·

#12

--, (2ND NULL POSITION ) OUTPUT VOLTAGE 180" OUT OF PHASE WITH INPUT VOLTAGE

.-------- ------------ ---------------- ------ - -------- ---- -------- ------------------ - -#1 (Repeat)

m

•w - ¿_e

400HZ

~- ·

(1ST NULL POSITION)

Figure 6-7-6. Voltage and phase relationship of an autopilot feed back synchro as it rotates 360°

6-22

1

Autopilot and A utoflight Systems PRIMARY ANO SECONOARY COIL WJNOINGS

LINEAR MOTJON

PUSH ROO (CONNECTEO TO THE MOVABLE SURFACE)

SOLIO METAL CORE CYLINOER

Figure 6-7- 7. Cut-a-way of a LVOT (li near voltage differential transducer)

PRIMARY WIRING

+-t-f - -----· rool l'lüll

SECONDARY WIRING

Figure 6-7-8. Wiring diag ram of a LVDT

AC synchros p rovid e excellen t feedback sign als for many autopilot systems. The phase shi ft principie, discussed abo ve, allows for acc urate measurements of even small control s urface m ovement. When placed in the null position, any movement clockw ise or counter clockwise is easy to m easure due to the phase shift and voltage change. The most accu rate m easurements u sing a synchro are th erefore obtained near the null positions. On most sys tem s, the synchro ro tor is connec ted to the servo output or control surface through a m echanical linkage, hence the synchro rotor moves in unison w ith the control surface.

Troubleshooting Synchro Systems Most synchro system s are fairly reliable. The electrical components a re simply w ire coils and ther efore seldom fail. The secondary pivot bearing can fail or become worn which causes inaccurate feedback s ig nals. Likewise, if the mech anica l linkage connecting the synchro becomes worn or binds during movem ent, inaccurate signals w ill result. The m echanical lin kage and pivoting secondary coi! are critica] and must have free movement. On m any systems, adjustment of the synch ro to the null p osition is critica! for proper operation. Many electric ser vos drive t he synchro to the null

position prior to engaging the servo clutch. This ensures the synch ro star ts in the null each time the autopilot is engaged . The electrical system of a synchro can be tested for the proper input voltage to the primary. In most cases, the input is 26 VAC 400 H z. The synchro primary and secondary coils can be tested for continuity and shorts to ground. When measuring continuity, it is critica! that the coi! r esistance be w ithin s pecifications. A change in resistance of just a few ohms can ereate inaccurate readings most li kely resulting from a breakdown of the coil's insulation. If an insulation breakdow n is suspected , be sure to monitor the system closely d uring the next se\·eral hours of operation.

Differential Transducers The differential transducer is typically used to provide a feedback signa! from hydraulic servos. There are two common ty pes of transducers used for autopilot feedback system s: the linear voltage differential transducer (LVDT and the rotary voltage differential transducer (RVDT). LVDTs and RVDTs produce a relatively weak electrical sig na] and are found on modem digital autoflight systems. Since they are u sed in conjunction w ith hydraulic servo system s, both LVDTs and RVDTs are ty pically fo und on transport category and high performance corporatetype aircraft. LVDTs and RVDTs are also fou nd in other non-autopilot systems to measure position or rate of mo tion. For example, the Airbus A-320 employs an LVDT on the turbine engine to measure stator vane p osition. The nose wheel steering system of the A-320 employs RVDTs to measure nose wheel position. The increasing popularity of LVDTs and RVDTs stem s from the simplicity of their d esign. As seen in Fig u re 6-7-7, the LVDT consists of a holJow metallic tube and a solid metal cylinder that is allowed to slide inside the tube. Around the tube are two electrical w indings, a primary

Autopilot and Autoflight Systems and secondary, similar toa transformer. A push rod is u sed to connect the solid m etal cylinder to the m ovable object that is being monitored. An LVDT or RVDT is a mutual inductive dev ice. The primary w inding is flanked by two second ary w indings as show n in Figure 6-7-8. The secondary w indings are wired to form a series opposing circu it. The primary receives an alternating current. This AC will induce a voltage into both secondary w indings. If the core m aterial is exactly centered, the output sig nals from the secondary w indings w ill cancel. As the core is d isplaced from center, the output signa! increases in amplitude. The phase of the AC output signa!, w ith respect to the input sign a!, is determined by th e direction of core displacement. Hence, the transducer can measure both d irection and magnitude of any movement. Figure 6-7-9 shows th e relation ship between core position and the secondary output signa!.

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Figure 6-7-9. The relationship between core position a nd LVDT output

from the LVDT is sent directly to the Flight Augmentation Computer (FAC).

Troubleshooting LVDTs and RVDTs Both LVDTs and RVDTs are relatively m aintenance free. The LVDT contains only one moving part, the core, w hich is typically supported so there is no contact between t he core and the coi! housing. The RVDT core is su pported by two bearings that have extremely long life due to the light load s on the input sh aft. Both LVDTs a nd RVDTs therefore have vir tu ally infinite mechanical life unless damaged by sorne externa! force.

Both LVDTs a nd RVDTs are always used in conjunction w ith sorne type of electronics circuitry. The circu it is u sed to interpre t the output sig n als of the LVDT or RVDT. On ma ny aircraft, the secondary's output signa! is sent to an LRU, which converts t he AC voltage and phase relationships into u sable data. Fig u re 6-711 sh ows a hydrau lically operated yaw damper servo containing an LVDT. The output sign a!

Electrically, the primary or secondary w inding of the transducer may fail due to an open or shorted circuit. In m any cases, a n ohmmeter can b e used to detect these failures. Simply disconnect the t ransdu cer from the aircraft wiring and perform a continuity test of the coils. Opens most often occur du e to induced stress on the w indings cau sed b y vibration, failed solder, or crimped connection. H owever, even

PRIMARY ANO SECONDARY WINDINGS

SECTION A-A Fig ure 6 -7-1O. Interna! co mponents of a typical RVDT

6-23

PRIMARY INPUT SIGNAL

An RVDT operates similarly toa LVDT except it is designed to detect rotation al movem ent. The RVDT contains a heart-shaped core material th at rotates w ithin a hollow tube (Figu re 6-7-10). As the core is rotated, it changes the output voltage and phase of the secondary. The rotational movement m easured by an RVDT is typically 120° or less, and the high est r esolution is obtained in the first 40° of rotation. An RVDT contains two bearing assemblies, required to support the input sh aft.

A

1

6-24

1

Autopilot and Autoflight Systems COMMAND SIGNAL FLIGHT AUGMENTATION COMPUTER (FAC

,.---f--HYDI~ LJLIC

PRESSURE INPUT .-+-+-HYI)RJ,ULIC PRESSURE RETUAA

FEEDBACK

Figure 6-7-11. An LVDT used in a yaw damper servo

these fa ilures are rare and the MTBF (mean time before failure) of a typical aircraft quality transducer is over one million hours. A nother means of troubleshooting an LVDT or RVDT is to measure the input and/or output voltage of the transducer. The input signa! must be an AC voltage within the limits established by the manufacturer. A high impedance voltmeter must be used for this test to ensure the meter does not distort the signa! being measured. The output sig na! from the secondary winding is determined by the position of the core. A dual channel oscilloscope can be used to show the voltage and phase relationship of the input a nd output signals. A voltmeter can be used to measure the transducer output voltage as the core changes position. This test is typically sufficient since it is virtually impossible to change the

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phase relationship of the o utput signa! once the transducer is installed correctly. As noted earlier, the output voltage of the secondary should be zero when the core is exactly centered (i.e., located at the null position). This condition rarely exists in the real world since the excitation voltage often conta ins high-leve: harmonics, which induce stray voltages into the secondary. Whenever measuring the output vol tage at the null position, remember that a small AC signa!, approximately 0.25 perc~ of maximum, may be acceptable. If the nuC position outpu t voltage exceeds th is amount be sure to check the purity of the input voltage to the pri mary. On transport categor y a ircraft, the troubleshooting process for LVDTs and RVDTs often becomes simplified through the use of built-in test equipment (BITE) or central maintenance systems. The integrated test equipment can be used to monitor the output of the transducer as the unit is m oved through its operating range. On many aircraft, a comparison is made of input and output signals to verify correct operation. For example, on the A-320 the Centralized Fault Display System (CFDS monitors engine stator vane position (Figure 6-7-12). The CFDS checks the input signa! to the stator vane actuator and compares that to the output signa! from the RVDT, which monitors the stator vane position. Using the CDFS, the technician can read command channel and monitor ch annel signals on a CRT display. The stator vane's position is measured in angular degrees. On this system, a 1° tolerance is allowable. The CFDS also performs continuous fault monitoring of the stator vane positioning system. If a fault were detected during flight.

Autopilot and Autoflight Systems the CFDS would take the necessary corrective actions and record the failure in memory for later recall. A transducer system found on the Boeing 767 is used to monitor aileron travel as specified by the autoflight computer. On this aircraft, the CMC (Central Maintenance Computer) monitors signals from three LVDTs. The technician can display transducer output on a CRT for comparison purposes. Whenever troubleshooting an LVDT or RVDT, always remember that these units are extremely reliable. In most cases, the associated w iring or electrical connectors are more likely to fail than the transducer itself. However, if a transducer is to be replaced in the field, take caution to en sure the proper installation of the new unit. Note any markings on the core and /or hou sing assembly. Be sure the core is installed in the correct configuration. Sorne LVDTs and RVDTs can only be replaced at an overhaul facility. In this case, if the transducer has failed, the en tire autopilot servo assembly must be replaced in the field. Whenever changing any transducer or servo assembly, always follow the manufacturer's instructions on installation and rigging very carefully. Sorne installations may also require a flight test of the autopilot system.

Tachometer Generators Tachometer generators, or tach generators, are often used in electric servo systems as rate sensors. The tach generator measures the rotational speed of the electric motor and provides feedback to the servo amplifier or autopilot computer. This feedback signal is typically used to regulate and limit motor speed. The tach generator consists of a permanent mag net and armature assembly. The generator spins in direct relationship to the servo motor. Hence, the electrical signa! produced by the generator is directly proportional to the motor's movement. A typical tach generator installation is shown in Figure 6-7-13. Tach generators can easily be tested for correct operation by measuring the output voltage as the generator spins. If the generator produces inadequate voltage it must be replaced.

Section 8

Collins APS-85 Autopilot System The APS-85 is a typical digital autopilot system found on high performance corporate-type aircraft. The system includes both autopilot and yaw dampening capabilities. The components

6-25

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include a mode select panel, (two panels are used for a dual system), a flight control computer, an autopilot panel, three primary servos, and three ser vo mounts (Figure 6-8-1). The APS-85 Flight Control Computer (FCC) is a dual ch annel system that provides redundancy for the autopilot. The FCC, located in the equipm ent rack, is cooled with forced-air. The three ser vos are mounted throughout the aircraft in appropriate locations for their respective control surfaces. The control panels are located on the flight deck and are accessible to both pilots.

Autopilot panel. The autopilot panel (APP) shown in Figure 6-8-2 contains the main controls for the system. The autopilot and yaw damper switches are gu ar ded levers and must be r aised to engage the respective systems. On sorne aircraft, the autopilot may be engaged independent of the yaw damper; on other aircraft, both systems must be engaged simultaneously. Prior to activating the autopilot/yaw damper, the FCC monitors the system for faults. If a fault is detected, the FCC will not actívate the autopilot/yaw damper. Whenever the autopilot/yaw damper is engaged, the appropriate message is displayed by EFIS. The APP pitch wheel is a spring-loaded rotary sw itch (Figur e 6-8-2). Moving the pitch wheel up or down modifies the vertical reference being flown by the autopilot. As a new vertical reference is entered, the value is displayed by EFIS. The turn knob is a bidirectional switch used to initiate a roll mode and define a given roll rate. The turn knob is inoperative while the autopilot is in the approach m ode. The autopilot transfer (AP XFR) switch is used on systems that employ a dual flight guidance system. The transfer switch is used to shift from the left to the right FCC. The turbulence (TURB) push-button switch is u sed to "soften" the ride when fly-

6-26

1

Autopilot and Autoflight Systems

MSP- 85 MOOE SELECT PANEL APP- 85 AUTOPILOT PANEL

MSP- 85 MOOE SELECT PANEL

SVO - 85 PRIMARY SERVO ANO SMT- 85 SERVO MOUNT

FCC- 85 /86 FLIGHT CONTROL COMPUTER

SVO - 85 PRIMARY SERVO ANO SMT - 85 SERVO MOUNT

Figure 6-8 -1. System components of the APS-85 autopilot

ing through rough air. When in the turbulence mode, the FCC lowers autopilot g ain sig nal s. This d egrades the inten sity of control surface m ovem ent.

Mode Select Panel The mode select panel (MSP) con sists of ten push -button switches u sed to control the various auto pilot modes (Fig ure 6-8-3). The variou s m o d e select push-buttons ar e som ewhat self d escriptive. The HOG (heading) switch commands the autopilot to steer a g iven heading. The 1/ 2 BANK m ode reduces all bank angles to approx imately 13.5 degrees. The 1/2 bank m od e is inoperative d u ring approach to land. The NAV (navig ation) m od e causes the autopilot to follow the n avig ation source currently displayed on EFIS. The APPR (app r oach) mode is u sed w hen the pilot w ishes to n avigate via a localizer and g lide slope sig na] durin g an appr oach to Jand.The NAV and APPR modes can be armed w ithout actually "capturing" t he mode. Capture of the m o de o ccurs only when valid n avigation sign a ls are available. If no val id n avigation sign a l (such as the localizer) is received the approach m ode will be armed (not captured). The autopilot w ill fl y the aircraft's last heading u n ti l cap t ure occu r s.

SVO - 85 PRIMARY SERVO ANO SMT- 85 SERVO MOUNT Courtesy of Rockwe/1/nternational, Collins Divisions

Vertical mod e switches include the: • CLIMB button is used to actívate a given climb rate (lAS or MACH ) according to FCC software. Three different rates can be selected (low, medium, or high speed) using the PERFORMANCE SELECT button. • ALT (altitude) button is used to maintain the current barometric altitude of the aircraft. • VNAV (vertical navig ation) mode is used to fly a vertical profile established by the flig ht management system. • OESCEND mode commands the autopilot to fly a preprogrammed d escent rate. The APP pitch w heel can be used to increase or d ecr ease that rate. • SPEEO mode will cause the autopilot to fly a given sp eed by adjusting aircraft pitch accordingly. Flight control computer. The flight control computer (FCC) is a dual channel unit d esigneC. to receive input data, process the information. and send the appropriate outputs to the autopilot ser vos and the electronic flight instrumertL system. While the autopilot is engaged, the FCC cont rols aircraft attitude throug h the control su rface serve s. With the au topilot engageC.

Autopilot and Autojlight Systems

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n u ~ Courtesy of Rockwelllnternational, Collins Divisions

overpower the motor in the event of a servo runaway.

Flight Control Computer Interface Servos. The APS-85 u ses three electrically actuated servos. Each servo is equipped w ith a n engage/ disengage clutch that allows for quick respon se time of the servo mechanism. The ser vos also employ a slip clutch that is used as a backup for ma nual override. The APS-85 is designed to interface with the aircraft's trim motor assembly and therefore the aircraft controls do not require additional trim ser vos.

Theory of Operation Refer to the block diagram in Figure 6-8-4 of the APS-85 system during the follow ing discussion on theory of operation. The FCC contains two channels (A and B), which receive identical inputs for data processing. The dual channels share the same FCC housing, yet operate completely independent. The FCCs perform system monitoring to ensure autopilot reliability and present all diagnostic data through the aircraft's EFIS displays. The FCC outputs control servo operation and flight director displays on the EADI. A voter circuit is contained within each channel of the FCC (Figure 6-8-5). The voter circuit determines which channel calls for the least servo movement and sends that signa! to the motor. A torque limiter is used as a current limiting device, which allows the pilot to manually

Inputs to the FCC include: l. Digital data in CSDB format from the two MSPs, the air data system, the Attitude Heading System (AHS), and EFIS. EFIS supplies all navigation inputs to the FCC. The FCC will also accept AHS data in ARINC 429 format.

2. Discrete inputs are received from the annunciator test sw itch, the autopilot disconnect switch, the go-around sw itch, the flaps switch, configuration strapping, options strapping (Lower center of Figure 6-8-4) as well as the pilot's and co-pilot's sync switches. 3. The APP sends analog data to the FCC for pitch and roll commands and autopilot/ yaw damper engage commands. 4. The three servo units send analog servo rate d ata to both channels of the FCC (Figure 6-8-5). The FCC outputs include: l. A two w ire 28 VDC an alog signa! to each servo motor. 2. A CSDB output is sent to EFIS w ith flight director information. The same data is sent to the air data system.

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6-27

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Attitude Heading System

The AHS-85 AHC uses a conventional flu x detector. The flux detector is needed to p rovide a magnetic heading reference to the system. Since each aircraft has slightly different magnetic character istics, a compensator unit must be used to correct for magnetic errors and flux detector misalignment. Remember, the compensator unit corrects for a specific aircrafts magnetic error, and therefore must s tay with the aircraft when changing other attitude heading system componen ts. The AHC contains a dual sensor assembly that houses two rotating wheels mounted at 90° angles from each other (Figure 6-8-8). The spinning wheels, which rotate at a constant 2,500 r.p.m ., contain fo ur p iezoelectric crystals called benders. One pa ir of benders measures acceleration, the other pair of benders measures rate changes (Figure 6-8-9). A convention al r otating mass gyro spins at 20,000 r.p. m. The slower rotational speed of 2,500 r.p.m. on the AH C

6-29

SERVO MOTORICLUTCH ASSEMBLY

ENGAGE CLUTCH

The APS-85 is designed to interface w ith a unique attitude heading reference system, the AHS-85. The AHS-85 measures angula r rates and accelerations along all three aircraft axes using piezoelectric sensors (Figure 6-8-6). The AHS-85 system employs an Attitude Heading Computer (AHC) containi ng the piezoelectric sensors. The piezoelectric sensors replace the accelerometers, gyros, and rate sensors fo und on conventiona l attitude heading systems (Figure 6-8-7).

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Figure 6-8-5. Diagram showing the connections be a two-channel FCC and a control surface servo

Figure 6-8-6. A piezoelectric sensor from the AHS-85 attitude heading system Cou rtesy of Rockwe/l lntemational, Collins Divisions

=-.:gu re 6-8-7. Attitude/ heading system components: (A) Early version systems with severa l indepen:;ent components-gyros, flux detector rate sensors, and accelerometers, (B) Newer version system -:ltaining all the necessary components in two units, the attitude/ heading computer and the flux ::=tector Courtesy of Rockwe/1 /ntemationa/, Collins Divisions

6-30

1

Autopilot and Autojlight Systems ing. Once the computer knows the magnetic heading, determined by the flu x detector, and the gravitational reference, determined by the rotating sensors, any changes in ang ular rate o r acceleration can be easily conver ted into changes in aircr aft position. These changes in position are transmitted to the autopilot FCC. Like the inertial reference system discussed earlier, the AHS must be initialized prior to use. This process requires that the aircraft be parked in an area free of magnetic interference. This allows for proper flux detector operation. The aircraft m ust also remain still during the initialization process. This system also has the capability to initialize d uring smooth, straight and level flight.

Figure 6-8-8. A dua l sensor assembly Courtesy of Rockwe/1 lnternotionol, Coffins Divisions

Inspection and Maintenance sensor assembly makes this unit much more reliable than a conventional gyro. When pressure is applied to a piezoelectric material, a voltage is produced. As the rotating piezoelectric crystals are subject to an acceleration or rate ch ange, the materials bend. Bending the material applies a pressure to the crystals; therefore, the benders produce a voltage. The direction in which the crystal bends, determines the polarity of the output voltage (Figu re 6-8-10). The rotating wheel contains a timing mark. This mark provides a reference point for AHC. The rotating wheels also contain two transformer primary coils. These transfor mers are u sed to induce the sig na! from the benders to the stationary portion of the sen sor assembly. It is important that the Collins AHC-85 be

oriented in the correct position for the system to oper ate properly. AHC configuration strapping is used to tell the computer which direction within the aircraft the AHC is fac-

The APS-85 requires very little routine maintenance. Operational tests should be performed in accordance with the aircraft's approved maintenance schedule. The servos are time-limited components and r equire regular maintenance. Every major aircraft overhaul or every 10,000 flight hours, the main control surface servos and servo mounts should be inspected by an authorized repair facil ity. It is recommended that any trim servos be inspected in as little as 1,000 flight hours. On the aircraft, the servo capstan should be inspected for wear, security, and proper cable alig nment. The servo unit should be operated through its entire range and observed. If a cable binding or fatigue is present, the problem must be corrected. If the servo unit makes a grinding or rubbing sound, the servo and servo mount should be removed, inspected, and repaired. At regular intervals, the ser vo slip clutch must be tested. The test procedures are outlined in the ser vice manual.

Acceleration Senslng Plezo Elements

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Figure 6-8-9. Sensor wheel assembly: (A) Diagram showing piezoelectric crystals (benders), (B) An expanded view of the sensor wheel Cour tesy of Rockwe/1/nternotiono/, Co/lins Divisions

Autopilot and Autoflight Systems

1

6-31

Troubleshooting Procedures The APS-85 is designed to operate in conjunction w ith the Collins electronic flight instrument system (EFIS). This allows the autopilot diagnostics to be displayed on the EFIS primary fligh t display or multifunction d isplay. The APS-85 diagnostics operate in three modes: input, report, and output mode. The input mode displays various input parameters that repor t to the FCC. The input mode can therefore be used to determine the operational outputs of systems, such as AH S, air data, and navigational aids. The report mode presents data on various systems that are monitored by the FCC. If a fault flag is d isplayed during fligh t, the report mode should be accessed prior to turning off electrical power. The output mode is used to display FCC software outp uts. In general, the report mode is used for primary troubleshooting, and the input/ output modes are used for detailed fault isolation. The pilot(s) of any aircraft using a n APS-85 should be m ade aware of the autopilot diagnostics. In the event of an autopilot problem, the pilot should enter the diagnostics mode on EFIS and record all fault codes prior to power shut down. This will help the troubleshooting process since the report codes give technicians important diagnostics d ata. The report codes are d ecoded using the autopilot maintenance manual. Report mode data includes categories such as REPAJR CODE (general fault d ata), AP DIS CODE (faults causing an autopilot disengage), STEER CODE (flight director steer ing faults), and RAM ERRORS (FCC memory fau lts). Diagnostic codes of 000000 indicate a system w ith no faults detected. Other fa u 1t codes must be decoded u sing the autopilot ma intenance manual. The diagnos tic procedures for using the Collins EFIS-85 were explained in Chapter 3 of thi s text.

Autopilot and Flight Director Problems Whenever troubleshooting any autopilot system, it is sometimes difficult to determine if the fault lies in the autopilot or the flight director. Keep in mind th at the: l. Autopilot computer typically controls the

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· ~·~· PORTION OF WAVE CREATED PORTION OF WAVE CREAT ED BY CRYSTALS MOVING THROUGH POSITION #1

BY CRYSTALS MOVING THROUGH POSITION #2

Fig ure 6-8-10. Diagram showing the voltage produced by bending crystal sensors: (A) As the crystal sensors rotate from straight down through position #1 to straight up the positive portian of the sine wave is produced, (B) As the crystal sensors rotate from full up through position #2 to straight down the negative portian of the sine wave is produced, (C) The sine wave is produced as the crystal rotates and the aircraft is under an acceleration force. The am plitude of the sine wave changes proportionally to the acceleration force of the aircraft. u sing the appropriate test sw itch . If an EFIS is u sed, check for an EFIS fault using the builtin diagnostics. lf either the electromechanical instru ment or EFIS show faults, the defect is in the indicator and the autopilot is most likely OK. If the indicator tests OK, su spect an autopilot problem. In general, w hen the autopilot is engaged, the flight director and autopilot functions are isolated . This is the best time to troubleshoot the two subsystems. If the pilot commands a left turn and the flight director respond s, but the autopilot does not, the fault is most likely in an autopilot component, perhaps a defective aileron servo. If the pilot commands the same left turn and the autopilot "flies" the aircraft into the turn, but the flight director does not respond, the fa ult is in the flight director, perhaps a fault in the ADI or EFIS interface. If both the fl ight director and autopilot fail to respond, the problem is in a component common to both subsystems, perhaps a defective APP or FCC.

flight d irector indications, and 2. Autopilot and the flight director typically receive different inputs from the autoflight computer. If the flight director indicator is an electromechanical type, test the indicator if possible,

Troubleshooting: Helpful Hints The following are severa! troubl esh ootin g techniques that m ay help isolate faults on the APS-85, as well as other autopilot systems.

6-32

1

Autopilot and Autoflight Systems

FLUX DETECTOR

FCC

AHC PROCESSING SOFTWARE

f ATTITUDE SENSOR ASSEMBLY

EFI

AHC ATTITUDE HEADING SYSTEM

Figure 6-8-11. Block diagram of an attitude heading system l. Many autopilot problems come from the

subsystems that feed the autopilot computer. One of the most complex and frequently failed subsystems is the attitude heading system. This is especially true for attitude heading systems that employ rotating mass gyr os. The AHS-85 contains both the flux detector for heading information, and the attitude sensors for attitude data. As shown in Figure 6-8-11, the attitude sensors are par t of the Attitude Heading Computer (AHC). Attitude and heading data are each processed and distributed through the AHC, which means: • If the EFIS displays a heading flag (HDG), the flu x detector is most likely at fault • If the EFIS displays an attitude flag (ATT), the attitude sensors are at fault

and the AHC must be replaced • And if both flags (ATT and HDG) are displayed, the AHC pr ocessor software is faulty and the AHC should be replaced. A lways become familiar with any subsystems that feed the autopilot, as this w ill aid you in troubleshooting. 2. If a YELLOW message such as HDG (heading) appears on EFIS, this normally means a dual system disagreement. That is, two redundant subsystems that feed the autopilot are not transmitting the same data. In this case, determine which unit is faulty by operating the s ubsystems independently. lndependent operation can be done through reversionary switching or by opening circuit breaker(s) to one of the subsystems. 3. Any time a heading inaccuracy problem occurs, consider that one of the fluxgates may be too close to a metal object. Move the aircraft and see if the problem corrects

itself. If a heading disagree fault is stored in the diagnostics memor y, check the time of occurrence. If the fa ult occurred shortly after star ting the engines, the fa ult is most likely caused by the aircraft taxiing too close to a metal structure. 4. While operating in au topilot mode, if the aircraft consisten tly changes altitude while in a banked turn, the fault is most likely a misalignment of the attitude head ing system. It is very important that the AHS-85 mounting tray be aligned correctly. An align ment fixture can be used to verify alignment of the AHS mounting tray. The tray can be shimmed to adjust the alignment if necessary. On any attitude heading system, if a misalign ment occu rs, inspect the mounting structure for cracks, bends, or looseness. 5. Any autopilot system is on ly as good as the related control surface elements. If the control surfaces are improperly installed, loose, or poorly bala nced, the autopilot will most likely be unable to hold a steady attitude. If the control surface cables are too loose, the aircraft w ill oscillate or porpoise while in the autopilot mode. This w ill be especially evident at capture of a given attitude or heading. The Phenom VLJ Autoflight System. One of the most modern aircraft designs at the time this text was wr itten is the category of aircraft known as ver y light jets (VLJ). These ai rcraft are typically constructed using composite materials in order to save weigh t and employ advanced integrated avionics which would typically include automatic flight control systems. The Embraer Phenom VLJ, introduced in 2008, has a capacity of fo ur to six passengers, can be flow n with one pilot, and has been well accepted by the industr y. Since the Phenom em ploys an avionics package constructed by

Autopilot and Autoflight Systems

PFD DISPLAY OF FUGHT DIRECTOR ANO AFC5 MODE ANNUNC/AT/ONS

AIR DATA COMPUTER 2 AIRSPEED ALTITUDE VERTICAL

INTEGRATED AVIONICS UNIT VHFCOM VHF NAV/LOC GPS GUDESCOPE

AHRS 2 ATTITUDE PITCH RATE OF IRNI===>J 5UP/5KID

MAGNETOMETER HEADING

AFCS MODE LOGIC FUGHT DIRECTOR (OMPUTATIONS SERVO MANAGEMENT

GIA2

2

Figure 6-8-12. IAUs process data from aircraft systems and pilot commands for control of the pitch, roll, and yaw serves Garmin International, the autoflight functions are similar to many small high performance aircraft also using Garmin avionics. The Phenom employs a Carmín Prodigy system similar to the Garmin G-1000 fou nd in the Cessna Mustang and other aircraft. As discussed earlier in this text, the Garmin integrated avionics system incorporates two PFDs and one MFD. These display units contain the circuitry for various software functions that deliver information to/ from the Phenom autoflight system. The two main processors in the aircraft are called the Integrated Avionics Units (also known as the GIA, Garmin

Integrated Avionics). These computers receive and process a variety of the data from aircraft systems as well as pilot commands from the fligh t deck. (Figure 6-8-12) The IAUs send information to the control su rface serve s for control of pitch, roll, and yaw. The servo actuators in the Phenom employ DC electric motors connected through various cables and mechanical systems to move the control surface. There is also a dedicated electrically operated pitch trip servo that moves a trim tab nota main control surface. The autoflight system found on the Phenom is called the Flight Guidance and Control System

1

6-33

6-34

1

Autopilot and Autoflight Systems HSDB

ARINC •29 AHRS2

AHRS1 ARINC -429

HSDB AOC1 ARINC 429

ARINC 4129 AOC2

HSDB

HSOB

INTEGRATED AVIONICS UNIT 1 (GlA 1)

~m

INTEGRATED AVIONICS UNIT 2 (GIA 2)

~~

C..\úro'MAric-¡:ü"Gfir coÑrRo_L_sv5T"E'Mf;uN"é.¡:,0-Ñs -:

AHRS1 ARINC 429 AOC 1 ARINC 429

: ¡

.

: AUTOMATIC FLIGHT CONTROL SYSTEM FUNCTIONS

: 1FLIGHT DIRECTOR 1 1

,._¡..:.:.:.._:..:,

¡ ' _

,

NORMAL PrTCH TRIM CHANNEL

'

IW
W.TCH

------

------

'' '' r-::---=-=t--.;'

r-::-::--:-::-::-:-::::--:--o--,

1FUGHT DIRECTOR 21

ARINC 429 AHR$2

NORMAL PITCH TRIM CHANNEL

........_

''

AUT~TIC

~----------------- ------

:

ARINC 429 AOC2

AUTOI-'ATIC W!TCH

'' ''

'

_____ :'

Figure 6-8-13. The Phenom Flight Guidance Control System interface connections (FGCS) and is divided into four major functions: l. Flight di rector (FD)

2. Automatic pilot (AP)

Courtesy of Embraer

the GIAs via a h igh-speed d igital bus (HSDB). The HSDB is a Garmin proprietary bus used on the G-1000 and similar Garmin systems. The GIA software performs the automatic flight control functions and outputs data to the appropriate servo actuators via a RS-485 bus.

3. Yaw damper/ turn coordinator (YD) 4. Automatic pitch control. The two GIAs compute the flight director and automatic pitch command functions. The three servo units each contain processing soft ware respon sible for automatic pilot and yaw dampening. Figure 6-8-13 shows the various interface connections of the FGCS on the Phcnom. The number 1 and 2 GIAs each contain identical software for flight director, yaw damper, and automatic pitch trim functions. Only one GIA performs FGCS calculations depending on the pilot's selection; the other GIA is ready in standby mode. The g uidance panel, located top left of the diagram contains most of the flight deck controls needed for the FGCS. According to pilot commands, the guidance panel (GP) sends RS-232 data to the PFD and MFD; these units process, convert, and send the GP data to

The servo units on this aircraft receive inputs from the GIAs and the pilot and co-pilot's control wheel autopi lot and quick d isconnect switches. These are considered "smart servas" since they contain software circuitry and process the incoming information prior to taking any servo action. Each servo contains two RS-485 transceivers and two processor circuits, providing redundancy. The servo processors exchange data, perform validity checks, and then control the servo motors as needed. The processors also transmit motor speed, torque, current, and voltage va lues to the GIAs as a feedback signal. Air data, attitude, and heading information is created by the ADC (air data computers) and the AHRS (attitude heading reference system). This information is sent directly to both PFDs and to GIAs vía an ARINC 429 data bus. Using independent busses provides redundancy and allows for validity checks. The ADC and AHRS

Autopilot and Autojlight Systems

1

6-35

information is processed by the GIA and used for automatic flig h t control system functions. Power to the FGCS comes from multiple busses to provide redundancy for the flight director function of the system. Remember the fligh t director provid es a visual reference to the pilot for aircraft guidance. The FD is therefore more critica! than the autopilot function and should be the least li kely to lose power. The servos and A HRS number 2 each receive power from DC bus number 2 only. If bus number 2 shou ld fail, the autopilot and yaw damper functions are inoperative. Since the GIAs receive power from both OC bus number 1 and 2, the GIA can continue to provide FO functions in the event of a OC b us failure.

ZONES

141 142

The Phenom flight controls and servos. Since this is a small light aircraft, the flig ht controls are r elatively simple. This aircraft employs the traditional connection between the flig ht deck controls and the actual control surfaces. The system incorporates push rod s, torque tube, cables, and bell cranks, as well as the three electromechan ical servos for autopilot and trim functions. As seen in Fig ure 6-814 each servo assembly consists of an electric motor, control circuitry, and the capstan dr ive. The cap stan employs a slip clutch common to most electromechan ical servos. The slip clutch is used for manual pilot override of the servo, if necessary, during a malfunction. The servos r eceive data packets from the GIAs through an RS-485 d ata bus. The data packets contain FD commands, attitude, rate information, accelerom eter data, and AHRS and ADC outputs. The ser vos also receive discrete information from the pilot's and co-pilot's disconnect switches. The servo assembly contains a solenoid activated dri ve w hich w ill au tomatically engage/d isengage the servo according to command. During autopilot functions, the servo force is transferred from the capstan drive through a stainless steel cable to the primar y control rigg ing. The contr ol su rface moves according ly as the servo rotates.

The Phenom flight director. The fligh t director function of Phenom employs two independent system s each located in the number 1 and 2 CIA. The FD function calculates pitch / roll commands and displays that data on the PFDs or MFD. The displayed data is used by the pilots for manual flig ht operations. Only one flight director function is operational at any g iven time; this d epends on which GIA is selected. Whenever the autopilot is engaged, all FD commands displayed on the PFDs should correspond to an asso ciated autopilot action. The FD software receives a variety of system and control panel inp uts, perform s logic cal-

A UTOPILOT AILERON SERVO

Figure 6-8 -14. A typical electromechanical servo as installed on t he Phenom Courtew of Embraer aileron flight controls

6-36

1

Autopilot and Autoflight Systems manually override the system using the slip clutch within the servo capstan.

Break-by-wire. The Phenom employs a brakeby-wire system activated through the traditional fight deck rudder pedals. When the brake pedal is pressed a signa! is sent to an electrically operated hydraulic valve which meters pressure to the brake actuators. As the pedal is pressed, a spring is compressed which provides a feedback force to the pilot. Nose wheel steering is mechanical on this aircraft, and sharp turns during ground m aneuvers can be m ade by applying asymmetr ic pressure on the brake pedals. This applies mor e braking force on one wheel and moves the nose wheel beyond its normal travel. Figure 6-8-15. An autopilot disconnect switch mounted on t he pilot's control yoke

culations as well as validity tests and health monitoring. The flig ht director has the capability to monitor and follow data from various navigation sources like GPS or VOR signals. The GIAs containing the FD func tions monitor the available inputs and if d ata is considered not available or invalid the GIA will ter mínate FD operations and the appropriate flag (red X) will appear on the PFD. The fligh t director can follow various modes of operation according to pilot selection, such as, approach (APP), navigation (NAV), and heading (HDG). The mode commands are selected by the pilot using the g uidance panel.

The Phenom autopilot. The autopilot function follows many of the characteristics of the FD system previously discu ssed. The AP is designed to eliminate pilot workload by automatically flying the aircraft w ithin limited operational parameters. Both processors within the autopilot servos determine the pitch /roll commands, and then initiate the correct motor speed/torque to move the control surface. Of course as with all autopilot system s, the servos m onitor actual servo movement in order to calculate feedback values. The GIAs communicate directly w ith each servo in order to calculate the appropriate AP annunciations, alerts, and engage/disengage logic. Th e AP function is independent of the yaw ·damper system and each may operate separately. The AP engage/disengage signals typically initiates from pilot commands and are controlled through the GIAs. The signals are sent to the three ser vo assemblies which then act accordingly. In order to ensure flig ht safety, the ser vos will autom atically disengage if the servo logic cir cuitry detects an interna! fault. The control wheel disconnect switch for the Phenom is shown in Figure 6-8-15. Of course if the servo d r ives fail to disengage, the pilot can always

Maintenance and troubleshooting the FGCS. Much of the maintenance for the flight guidance and control system will be in the form of software updates and reinstallations. Updates w ill be needed at regular intervals and software may need to be reinstalled if components are removed and replaced. As described earlier in this text, the configuration changes are made using an SO card and a series of installation step s. Be sure to consult the aircraft manual for complete instructions. The troubleshooting process as described in Chapter 5 of this text is also well explained in the aircraft manual and typically involves access to maintenance pages through the PFD. The systems configu ration page will display red, gr een, or black indications to show system /component status. The manu al is then referenced to determine the corrective action. Periodic maintenance of the flight g uidance system includes sorne of the items below. In general only the electromechanical ser vo assemblies and pitot/static system require regular testing or ser vice. Since this text is for training purposes only, be sure to refer to the current approved mainten ance data prior to any testing, maintenance, or service activities. The items related to the pitot/static system include: l. Perform a pitot/static leak test to ensure system integrity. Th is test mu st be performed every 24 calendar months in accordance w ith FAR 91.411 Part 43, Appendix E. Th is test is performed to ensure the pitot/static system is not leaking and the VSI (vertical speed indicator) and the altimeter are w ithin acceptable limits. These instruments are critica! since this aircraft would typically fly in airspace requiring reduced vertical separation minimums (RVSM). RSVM aircraft require an accuracy inspection every 12 calendar months.

Autopilot and Autoflight Systems 2. The Mode S transponder must be tested in accordance with FAR 91.411 and 91.413, Append ix F when flying in United States airspace. This test is performed to ensure correct operation of the aircraft's altitude reporting transponder. 3. The Garmin magnetometer, which is u sed to determine aircraft compass heading, must be updated every five years. This is done to ensure the magnetic compass system (the Garmin GRS77) software u ses the latest version of the earth's magnetic field model.

6 -37

AIRCRAFT AUTOFLIGHT SYSTEM (1) INERTIAL REFERENCE SYSTEM (IRS) .._ POWER INPUTS AIR DATA..-.. (2) FLIGHT MANAGEMENT SYSTEM (FMS) (3) AUTOPILOT FLIGHT DIRECTOR SYSTEM (AFDS) .._ MISCELLANEOUS AIC SYSTEMS ELECTRONIC_.. (4) YAW DAMPER SYSTEM (Y/D) _.. DATA LOADER INSTRUMENT +-+ MISCELLANEOUS CONTROL SYSTEM PANELS CENTRAL ........... MAINTENANCE ~-.-J---.¡-----,Lr-----' COMPUTER SYSTEM CON TROL SURFACE. YAW DAMPER, (CMCS) NAVIGATION ANO AUTO THOTTLE SERVOS RADIOS

<&9~J~~L UNITS

Figure 6-9-1. The four subsystems of the B-747-400 Autoflight system

Inspection and maintenance items related to the servo assemblies include: l. Visual inspection of the servo assemblies ever y 1,000 hours and /or every annual inspection oras stipulated in the progressive inspection schedule. This inspection should include all electrical w iring and connectors, support structure and mounting hardware for cracks, deformation, excess ware, or dirt.

CMC'S

2. Clean and grease the servo assembly in accordance with the aircraft maintenance manual every 1,000 hours or three years. 3. Visual inspections of the servo assembly slip clutch every 500 hours or one year. 4. Visual inspections of each servo cable for corrosion, chaffing, fraying, excess ware or other defects during annual or progressive inspections. 5. Check the tension of each servo cable and adjust as needed according to procedures outlined in the maintenance manual during annual or progressive inspections.

Figure 6-9-2. IRS interface diagram

3. Autopilot flight director system 4. Yaw d amper system The autoflight system also receives data from and communicates with a variety of other aircraft systems. The follow ing discussion on the B-747-400 autoflight system will include an overview of the four major systems listed above.

Inertial Reference System Section 9

The Boeing 747-400 Autoflight System A modern autoflight system operates using microprocessor technologies, and communicates with a variety of aircraft systems via digital data busses. The Boeing 747-400 autoflight system is typical of a fully integrated digital system found on modern transport category aircraft. As seen in Figure 6-9-1, the B-747-400 autoflight system is comprised of four major subsystems:

The B-747-400 inertial reference system (IRS) is used to provide vertical and horizontal navigation, attitude information, acceleration, and speed data to a variety of airo·aft systems. The IRS consists of one mode select unit located on the flight deck and three inertial reference units (IRU) located in the ma in equipment center. As seen in Figure 6-9-2, each IRU interfaces with the air data computers, control display unit/ flight management system, central maintenance computer system (CMCS), integrated display system (IDS), and the various IRS data users.

Controls l. Inertial reference system

2. Flight management system

The IRS mode select unit (MSU) is used to select one of the three IRS operating modes:

Courtesy of Northwest Airlines, /nc.

6-38

1

Autopílot and Autoflight Systems

FLTOIR L

FLT OIR R

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IRSSOURCE SELECT SWITCH ES

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AIROATA R

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align, navigation, or attitude. The MSU can also be used to turn off each individual IRU (Figure 6-9-2). While in the align mode, the IRU performs the alignment procedures to determine the aircraft's local vertical, heading, and present position. While in the navigation mode, the IRU w ill provide: attitude data, accelerations, heading data, horizontal and vertical velocities, wind speed and direction, latitude and longitude, ground speed, and iner tial altitude. If the mode selection switch is moved from OFF to the NA~ the IRU will perform the alignment procedures befa re operating in the navigation mode. The navigation mode is used during normal flight con. figurations. The IRS attitude mode is a backup mode used in the event the navigation mode fails. The attitude mode provides attitude data, heading information, accelerations, and vertical speed. There are two source select panels located on the flight deck that are used to choose wh ich IRS w ill provide data to the electronic flight instruments (Figu re 6-9-3). The captain and first officer can each selecta different IRS source for their resp ective electronic flight displays. The first officer's TRS select switch also controls the source for the standby RMI.

Operation

FIRST OFFICER

CAPTA! N

Figure 6-9-3. B-747-400 IRS source select panels

REFEREN eE AIRPORT IOENT

POS INIT

Courtesy of Northwest Airlines, lnc.

112 LASTPOS

N40• 38.0 W073• 46.4

8§F_ . GATE

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rrnrn.o OJIJ·rn. GMT SeRATe H PAO

UNE SELEeT KEYS

MESSAGE UGHT {WHITE)

e OU (TYP)

Figure 6-9-4. Control display unit showing the IRS disp lay during initialization procedures Courtesy of Northwest Airlines, tnc.

At power-up, the IRU computer performs a BITE test that verifies the health of the system. The test monitors interna! circuitry and the power supply switch-over capabilities. Each IRS has the ability to automatically switch fro m 115 VAC to 28 VDC power in the event of a bus failure. If the IRU passes the powerup test, the unit moves in to the eight-minute initia lization process. At this time the pilot must enter the aircraft position (latitude and longitude) using the control display unit. Figure 6-9-4 shows the typical display during latitude/ longitude data entry. During initialization, the white memo message IRS ALIGN MODE L/C/R will be displayed on the main EICAS display. If the alignment process is disturbed by excessive aircraft motion, the main EICAS display will show an amber advisory IRS MOTIOl\T (Figure 6-9-5). This message will be displayed until 30 seconds after the motion stops. At this time, the corresponding w hite memo message will be removed from the display. The IRU will automatically continue the alignment procedure when the motion stops. After alignment is complete, the three IRUs compare position data to ensu re accuracy. If a miscom pare message is displayed by EICAS, the alignment procedure should be repeated. Once alignment is complete, the IRS is ready for operation.

Autopilot and Autoflight Systems

AMBER ADVISORY MESSAGE

IRS MOTION

D D D D

- 1111 D

GROUND TESTS 1/1 34 INERTIAL REFERENCE - CID THE "IRS LEFr E ICAS AOV1SORY MESSAGE APPEAR?

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D D D D D D

D D D D D D

1

6-39

GROUND TESTS XIX IRU·l ' FAULr ANALOG DISCRETE-EIU'S INTERFACE FAIL MSG:347 17

ATA:34- 21

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D D D D D D

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MAIN EICAS DISPLAY (P2)

Fig ure 6-9-5 . IRS display caused by excessive motion during alignment Courtesy of Northwest Airlines, lnc.

D D D D D D

GROUND TESTS 111 34 INERTIAL REFERENCE

D

FAIL>

D D D D

Figure 6-9-6. Typical sequence of displays for an IR$ ground test Courtesy of Northwest Airlines, lnc.

Maintenance and Troubleshooting Whenever troubleshooting the IRSs, remember each IRU (left/right/ center) receives power from different sources. Also, each IRU must be supplied with 115 V 28 VAC power and 28VDC power supplied from the APU hot battery bus. If the APU battery is below 18 volts, or is removed from the aircraft, the IRUs will fail the power-up test and w ill not func tion. Each of the three IRSs is monitored by the aircraft's central mainten ance com puter system (CMCS). If a fa ult occu rs d uring operation, the CMCS will cause the ap propriate message to be d isplayed by EICAS. To veri fy the fault and determine the suggested repair, the CMCS can be accessed through the control display u ni t. A ground test can be performed on each IRU using the CMCS. The test is accessed through the ground test menu of the CMCS. The technician should select the appropriate IRU for testing and follow the test preconditions. Pressing the line select key adjacent to START TEST initiates the test process. If the ground test is passed, the CMCS displays a question asking if the advisory message IRS LEFT was displayed by EICAS during the test. The technician should answer the question accordingly. If yes is selected, the message PASS appears on the CMCS display. If no, the appropriate ground test message is displayed as shown in Fig ure 6-9-6. As mentioned earlier, du ring power-up of the IRS, an IRU BITE test is automatically conducted. The IRU BITE can also be activated by the interface test switch located on the face of the IRU. If the IRU fails its interna! BITE test,

the fa ult ball will be visible on the face of the IRU (Figure 6-9-7). During removal and installation of the IRU, be sure to handle the u nit gently. Also, the IRU must be installed w ith precise alignment if accurate output data is to be obtained. Each IRU mounting rack contains an alignment pin, which must fit accurately into the IRU alignment hole for proper installation.

Flight Management System The flight management system (FMS) is a computer-based system that reduces pilot workload by providing automatic radio tuning, lateral and vertical navigation, thrust management, and the display of flight plan maps. Automatic radio tuning is performed IDENTIFICATION PLATE

HANDLE

FAULT BALL BLACK-NO FAULT YELLOW-INTERNAL FAULT INTERFERENCE TEST SWITCH

Figure 6-9-7. Diagram of the IRS inertial reference un it Courtesy of Nort hwest Airlines, lnc.

6-40

1

Autopilot and Autoflíght Systems

O

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by the FMS for all navigational aids used during normal flight. Radio tuning includes selection of the appropriate radio, tuning to the correct frequency, and selection of the correct course bear ing. Vertical and lateral navigational parameters are computed by the FMS and sent to the fl ight d irector and autopilot systems. Thr ust ma nagement automatically controls engine thrust as needed for a given flight cond ition. The fl ight plan map displays are constantly updated by the IRU computer according to the programmed flight plan . The flight plan map s are displayed on the EFIS CRTs. The flight management compu ter (FMC) is the main element that provides interfacing and data processing for the FMS

Controls

NUMERIC KEYS

ALPHA KEYS

Figure 6 -9-8 . Typical control display unit (CDU)

Courtesy of Northwest Airlines, lnc.

l..!!.:i!,,,,I--H/ 1 - - - AUTOTHROTT\.E DISCONNECT SWITCH (571

THRUST LEVER NO. 1 (NO. 4 SIMILARI

Figure 6-9-9. Diagram showing the autothrottle disconnect switches located in the number one and four thrust levers Courtesy of Northwest Airlines, lnc.

NAVSOURCE .....___ - - - - - SELECT SWITCH ----....

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CAPTAIN'S NAV SOURCE SELECT SWITCH

NAV

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The B-747-400 contains th ree control display units (CDU), which are u sed to enter data into the FMS a nd interface the FMS w ith other aircraft sensors and systems. The CDu keyboard contains fou r typ es of push-button switches: alpha numeric, mode, line select, and function (Figure 6-9-8). These switches are often referred to as Keys. The line select keys perfor m a specific funct ion according to the items displayed on the CRT; all other keys perfor m a given fu nction as labeled on the key. There are four annunciators, which illuminate to show specific messages related to the CDU. The display (DSPY) annunciator illuminates whenever the currently displayed page is not related to the active flight plan. The FAIL annunciator illuminates if the selected FMC fails. The message (MSG) illuminates if a message appears in the FMC scratch pad. The offset (OFST) light illuminates when navigating using an offset rou te. The annunciator lamps are accessed by removing two screws that hold the annunciator assembly to the CDU faceplate (Figure 6-9-8). The autothrottle (A/T) system is activated through the mode control panel, to be discussed later. The autothrottle disconnect switches are located in the number 1 and 4 thrust levers as shown in Figure 6-9-9. Two switches are activated by each disconnect lever to provide redundancy.

R

e

FMC MASTER SWITCH

00]

E

S E L E

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FIRST OFFICER'S NAV SOURCE SELECT SWITCH

EICAS CONTROL PANEUFMC

Figure 6-9-1O. The flight management control system navigatio n so urce Courtesy of Northwest Airlines, lnc. select switches

The FMC master switch is located on the EICAS control panel (Figure 6-9-10). The master switch selects w hich FMC (left or right) will control commands for: autopilot, autothrottle, and radio tu ning. The navigation source select switches control which FMC is u sed to drive EFIS displays. The navigation source select switches and the EICAS/FMC control panel are located on the instrument panel.

Autopilot and Autoflight Systems

Architecture The B-747-400 contains two complete flight management systems (right and left). The flight management computers (FMC) are located in the main equipment bay and perform all the necessary interface and data processing functions for the FMS. The flight crew selects one FMC as "master." The opposite side FMC opera tes in hot standby in the event that the master FMC fails.

computer system (CMCS). The CMCS stores all fault data that can later be retrieved by the technician for analysis and fault isolation.

Autothrottle Architecture The autothrottle func tion of the FMS is regulated through the mode control panel (MCPt which is located in the center of the flight deck glare shield. As seen in Figure 6-9-12, the MCP communicates directly to the r ight/left FMC. The FMC then transmits control signals to the autothrottle servo. Whether the th rottles are moved manually or by ~utoth rottle, a feedback signa! is sent from the throttle resolver angle (TRA) transducers to the EECs. The EECs send data to the engine fue! control units (FCU), which provide "coarse" adjustments of engine thrust. The FMC provides "fine" adjustment of engine thrust.

An FMS interface diagram is shown in Figure 6-9-11. The FMC receives data from the control display unit and a variety of aircraft systems a nd sensors. A d ata loader is used to input preprogrammed navigational parameters, such as flight routes, way points, and airport data. Sorne of the FMC outputs are sent directly to the user, while sorne are sent via the FMC master relays. The use of relays allows either FMC to send critica! output information to four syste ms: the EECs, FCCs, MCP, and the n avigation radios. The FMC sends output data to the electronic engine controls (EEC), which provide the control signals to the autothrottle ser vo motor. The ser vo motor generators send a feedback signa! to each FMC. The flight control computers (FCC) receive FMC data for control of autopilot and flight director functions.

The FMC provides engine trimming commands. Trimming the engines is simply a fine thrust adjustment in order to precisely equalize the thrust of all four engines. The FMC receives engine thrust data from the EFIS/ EICAS interface units (EIU) and calculates the trim commands. The trim commands, along with air data computer (ADC) information, are sent to the EECs.

The FMC sends output data to the integrated display system (IDS). This data is used to d isplay FMS information on the PFD, NO, and EICAS. The data sent to the IDS is also u sed to communicate w ith the central maintenance

An autothrottle/ FMC interface diagram is shown in Figure 6-9-13. The autothrottle assembly contains both a servo motor and a tachometer generator. The tachometer generator sends a feedback signa! to each FMC. The FMCs send

DATA LOADER

• •

;~J.-0 H§!§!ll~ ~~~j;--~~=r-1 o

00E>a:JEEB

CONTROL DISPLAY UNIT

OTIIER SYSTE11S SENSOR S • IRS • ADC • NAV RADIOS • fQIS • AFOS • CLOCK

Fig ure 6-9-11. Block diagram of the flight management system

Courtesy of Northwest Airlines, lnc.

1

6 -41

6-42

1

Autopilot and Autojlight Systems VERT SPO

lt 1 1 1 1 tJ

F.O

ON

g

f~~ '-.,

OFF

UP

MCP

A LT

ffÍtl rlolo¡o!3

li:!!

F.O

ON

&\

1

o

OFF

11 AIT ARM

AIT ARM

TRIH

FCU/IIMU

Figure 6-9-12. Thrust management system interface diagram.

Courtesy of Northwest Airlines, lnc..

-

~ ~-

LEFT SERVO ORIVE FWD REV SERVO EXCITATION L

} AIT DISCONNECT 1 RESET

TllRUST LEV ER NO. 1

~ lli8

':

THRUST LEY ER

~¡::¡¡,---

r l- 1- 60 AROUND LEFT FMC TACII FEEDBACK

1•

~-

111

LO

~ } AIT DISCONNECT 1 RESET

TttRUST LEVER

SERVO EXCJTATION R RI6HT SERVO I>RIVE

~-

FWD

,.....---

~

THRUST LEVER NO. 3

60 AROUND Rl6HT FMC

Figure 6-9-13. Autothrottle/ FMC interface diagram

servo commands to the autothrottle servo motor consisting of a llSVAC excitation voltage anda 28 VDC forward and reverse signa!. The autothrottle disconnect and go-around signals are sent from the throttle lever switches to each FMC. The FMCs then senda discrete signa! to the autothrottle servo motor generator assembly to command go-around or disconnect.

REV AUTOTHROTTLE SERVOMOTOR GENERATOR Courtesy of Northwest Airlines, /nc.

FMS Power Inputs Power inputs to the FMC come from six different circuit breakers and five different p owe¡distribution busses. The 28 VDC busses 1 and 2 supply power for the autothrottle servos ané master relays 1 and 3. The captain's 115 VAC tra nsfer bus supplies power to the left side

Autopilot and Autoflight Systems FMC for interna! FMC functions. The first officer's 115 VAC transfer bus supplies autothrottle servo excitation power along w ith tachometer generator excitation. The 28 VDC battery bus powers the FMS warning circuits. The first officer's {F/0) 115 VAC transfer bus powers the interna! functions of the right FMC.

stored in the CMCS memor y can be accessed through the CMC existing faults or present leg faults page.

Autopilot Flight Director System

Maintenance and Troubleshooting

The B-747-400 autopilot flight d irector system (AFDS) receives inputs from various systems and sensors th roughout the aircraft, and provides steering commands for automatic and 1 or manual control. For manual steering, the flight director provi~es the interface between the AFCS (automatic flight control system) and the pilots. Du ring automatic steering, the aileron, elevator, and r udder servos provide an interface between the AFDS and the control surfaces. The autopilot is capable of pitch control to maintain a given airspeed, altitude, vertical speed, or ver tical navigation including glide slope. Roll commands can maintain a given heading, track, lateral navigation, or attitude including localizer. The autopilot yaw func tion provides control for adverse yaw, and crab angle.

The FMS continually monitors itself using BITE systems programmed into the FMC software. The BITE is initiated at every power-up of the FMC. The BITE can also be initiated through the central maintenance computer system or using the INITIATE TEST/LAMP TEST switch on the front of the FMC (Figure 6-9-14). During this 15 second test, the main and auxiliary EICAS, the PFD, and ND each present specific test messages. During the test, the master caution and warning lights and aural tones sound for a short period. On the FMC, the red FAIL lamp illuminates while the FMC test switch is held in, or at the end of the test if the FMC BITE fails. The TEST IN PROCESS light illuminates any ti me thc test is in progress. Two major subsystems of the FMS can be accessed through the central maintenance computer system (CMCS): the FMC and the FMC servo loop. Both of these systems can be accessed fro m either the right or left FMS. The CMCS tests for the FMS can only be performed on the ground since the FMS is inoperative during CMC interrogation. FMS fault data

Controls The mode control p anel (MCP) is the main interface between the fligh t crew and the AFDS. The mode control panel is located on the glare sh ield, cooled by forced air, and connected to the system th rough three connec-

> AIITOTHROT OISC > FMC LEFT > FMC MESSAGE

o o

o MAIN EICAS (P2) FAULT ANNUNCIATOR (RED LEO)

o

-

CAUTION

1CAUTION

AURAL WARNING

MASTER CAUTION LIGHTS

FMC L

T EST IN PROGRESS {YEL LOW LEO)

0 INITlATE TEST/ LAMPTEST

o

AUX EICAS (PS)

MAINTENANCE SELF-TEST SWITCH

o PFD (P1)

' MAP I FMC VOTEST OK

FLIGHT MA NAGEMENT COMPUTER -LEFT

/::¡

m ~

ND (P1 )

Figure 6-9-1 4. Flight ma nagement control system BITE test can be controlled by the maintenance selftest switch fou nd on the FMC. Courtesy ot Northwest Airlines, lnc.

1

6-43

6-44

1

Autopilot and Autoflight Systems

VERTICAL SPEEO SELECTION VERT SPO

ALTITUOE SELECTION

AUTOPILOT ENGAG E CONTROLS

FIRST OFFICERS FLIGHT DIRECTOR

ALT

11 1 1 1 1 JI l!tlr!o!o!olj

o'

tflt'-rwsl \jf. LJ t:j

UP

MCP FRONT VIEW AIR INLET

IOENTIFICATION PLATE

TEST CONNECTOR (2)

REAR CONNECTORS MCP REAR VIEW

AIR EXHAUST (2)

5 V AC BULBS (2)

TYPICAL LIGHTEO PUSHBUTTON SWITCH

Figure 6-9-15. Boeing 747-400 Autothrottle Flight Director System (AFDS) mode control panel Courtesy of Northwest Air/ines, tnc..

tor plugs located on the rear of the unit. In Figure 6-9-15, a lighted p ush button assembly is removed fro m the face of the unit for lamp replacement. Each lamp assembly contains four bu lbs, two powered by 5 VAC, and two powered by 28 VDC Refer to Figure 6-9-15 dur ing the MCP control explanation in this paragraph. The captain's fligh t director is activated by the toggle switch on the far left of the MCP; the first officer's flight

AUTOPILOT DISENGAGE SWITCH

AUTOPILOT DISENGAGE SWITCH

director toggle switch is located on the right of the panel. The autothrottle engage sw itch is located just right of the captain's flight director switch. Indicated airspeed (lAS), or mach speed, can be selected from the speed mode of the autothrottle function. Lateral n avigation (l NAV) or vertical navigation (V NAV) can be selected using the appropriate lighted push button switch. Pressing the flight leve! change (Fl CH) switch w ill engage both vertical and lateral navigation. The HDG control can be used to selecta given heading for the autopilot or flight director. Vertical speed is entered into the MCP using the vertical speed thumb wheeL A given altitude can be selected and displayed in the ALT w indow. The autopilot engage push buttons allow the pilot to select the left, center, or right FCC for command of autopilot/ flight director function s. An autopilot disengage switch is located on both the captain's and firs t officer's control wheel. These switches are removed by a screw located on the front of the switch plate (Fig ure 6-9-16). The switch wiring is fed through the control w heel toa terminal block. The autopilot go-around switches are located on the number 2 and 3 thrust levers (Figure 6-9-17).

Architecture

Figure 6-9-16. Autopilot disengage switch

Courtesy of Northwest Airlines, lnc.

The three FCCs interpret data and provide the necessary calculations for the autopilot and fJigh t director functions. The pilot

Autopilot and Autoflight Systems selects inputs to the FCC through the MCP, the heading reference switches, the disengage switches, and go-around switches (Figure 6-918). The FCC receives three types of system inputs: navigational, a irplane configuration, and triple redundant sen sors. Navigational inputs are provided by the FMC and ADC. Airplane configuration sensors monitor items necessary for autoflight, such as hydraulic statu s and flap position. The triple redundant inputs are those needed for autoland functions. Triple redundant sensors include: ILS, IRU, and radio altimeter data. The three FCCs each control a separate servo, one each for the ailerons, elevator, and rudder. The ser vos use electrical signals from the FCCs to control the flow of hydraulic flu id, which in turn controls the position of the related control surfaces. The FCC outputs display data to the EFIS/ EICAS interface units (EIU). As seen · in Figure 6-9-19, al! three FCCs send a parallel data signa! to each of the EIUs. The FCCs send a discrete warning signa) to the modularized avionics and wa rning electronic assembly (MAWEA) for annunciation of warning data. Caution information is sent from the FCCs to the three EIUs. The FCCs communicate to each other via a crosschannel data bus fo r exchange of health monitoring, and . to provide redundancy for servo engage data. The ability to cross ta lk between

Al}~RM

@

FJD

ON

g, OFF

SWITCH

Figure 6-9-17. Autopilot go-around switches FCCs improves system sa fety by allowing the comparison of information between computers. If any FCC detects a failed FCC or critica! system out of tolerance, the autoland capability will not be available.

~EJÉ]

OFF

~-0--Ej

"

F.O

ON

B

OFF

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•

RIGHT GO-AROUND SWITCHS6 (SS)

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S

IHTEGRATED ------------DISPLAY SYSTEH

SEHSORS

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• FHC AIRPLANE CONFIGURATIOH SEHSORS • SPD BRK HANDLE POSITION • HYORAULJC VA LlO • AlR/GROUND • FLAP/STAB CONTROL

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ELEVATOR SERVO

UNIT

TRIPLE REOUNOANT SENSORS • JLS • JRU • R~D ~LT STAB TRIH/RUOOER R~ TJO MODULES CENTRAL HAINTENANCE COHPUTERS BUS CONTROL UNITS HAWEA

AILERON SERVO ROLLOUT POIJER CONTROL PACKAGE

PILOT INPUTS HEADING REF SWITCII

6-45

LEFT ANO CENTER GO-AROUND SWITCH ES S2, S4 (S1 , S3)

IASIMACH

1 illJ 1 ~ SEL~/ lfiiiil ' ~ @ ti.AUlO-<

1

~

olSENGAGE SWITCH

=iJ

GA SIJITCHES

Figure 6-9-18. Interface diagram of the flight control computer (FCC) and various aircraft systems Courtesy of Northwest Airlines, lnc.

Courtesy of Northwest Airlines, lnc.

6-46

1

Autopilot and Autoflight Systems

A/P 1./ARN 2 NORMAL A/P 1./ARN 2 BATTERY A/P WARN 1 BATTERY

WARNING RESET WARNING LIGHT HAWEA

• • • • • •

A/P WARN 1 NORMAL A/P CAUTION ADVISORY HESSAGES STATUS HESSAGES AFDS MODE AUTOLAND STATUS ENGAGE STATUS FLIGIIT DIRECTOR COHHAN OS • FLlGIIT CREW SELECTED DATA

AURAL WARNING SPEAKERS

-

CAUTION AURAL A/ P WARNING

•

•

41 WARNING

EIU L

FCC L ¡--

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SAME AS FC C L

41 WARNING

~

SAHE AS EIU LEFT EIU

fCC C

r-

CAUTION}--F/ 0 MAS TER/ CAUT 1ON LI GIIT

¡.-

e

SAME AS EIU LEFT

¡-¡-----

SAME AS FCC L

CAUTION CAPT MAS TER/ CAUTION LIGIIT

CAUTION LIGHT CAUTION RESET :+-

¡._

EIU R

FCC R Figure 6-9-1 9. FCC/ EIU/ MAWEA interface diagram

Courtesy of Northwest Airlines, /nc.

28V OC BUS 3 28V~BUS2

4

MD&T

,----. MCP

FMC Mil STER RLY 2

fl

R FCC

lf

FMC MIISTER RLY 1

lf:®! J¡

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28V OC STBY BUS 116V A C BUS 3 28V OC BU$ .3

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2

{ lll TUNE INHIBIT LILS

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r;

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C FCC

SAMEAS CFCC

TEST IWHIBIT LRAO { ALT CAPT AIP OISENGAGE SW

R FCC

L IRU

cr

CAPTADC

GO.AROUNOSW AIRIGROUNO SYS CAPT F/0 SOURCE SEL SW FIO FIC SOURCE SEL SW HOGREFSW GISANTSW

26VAC A RM SOLENOIC ENG.A SOLENOI O SERVO LVOT SURFACE LV OT

l

L ELSV AIP SERVO

3

SAMEAS ELEV AIP SERVO

2

L CENTRAL LATERAL CONTROL PACKA GE

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LCMC {

HYORAULIC VALlO HANOLE POSITION FLAP PO SITIO N STAB POSITION BUS ISOLATE BUSSES ISOLATED

3

SAMEAS EI.EV AIP S ERVO

2

L ROLLOUT POWER CONTROL PACKAGE (A AUTOTRIM VALlO

GROUNCTEST

I.SRM ElU$(3) WA RNING WARNING

RCMC { d

2

L EIU

M.AWEA

LFCC

Figure 6 -9-20. Left FCC interface diagram, note t hat th e center and right FCCs receive similar inputs

Courtesyof Nor thwest Airlines, lnc..

Autopilot and Autoflight Systems Figure 6-9-20 shows the interface of the FCC and various aircraft systems. In the upperright portion of the diagram are the cross channel busses for communication between the left/ right/center FCCs. The data busses to the central maintenance computers (CMC) are shown in the lower left portion of the diagram. Discrete data, represented by a single line on the interface diagram, comes from a variety of other aircraft system s to the FCCs. In the top left portian of the interface diagram are the power inputs to the FCC and MCP. To operate this autoflight system, there are a total of nine different circuit breakers fed from seven different power distribution busses. Whenever troubleshooting the system, make sure power is ava ilable to all necessary circuits.

Maintenance and Troubleshooting The B-747-400 Autopilot Fl ight Director System (AFDS) contains BITE circuits that continuously monitor the health of the FCCs and related systems. The BITE circuits are located within each FCC and report all autoflight failures to the central maintenance computer. The flight crew is made aware of failures by a flag on the PFD or NO, a n EICAS m essage, and /or a discrete annunciator and audio tone. EICAS w ill always display a warning, caution, advisory, or status message for the various AFDS faults. The A/P DISCONNECT message is the only EICAS warning applicable to AFDS. Remember, warn ings are the most serious EICAS message and require immediate crew action . This message w ill display on EICAS for either a manual or automatic disconnect. In the event of a m anual (pilot activated) disconnect, the CMC w ill not store the message as a fault. Any fault that is sen sed b y the BITE circuitry is automatically r ecorded in the CMC nonvolatile m emory. The technician can access cu rrent or previous failures through the Existing Faults or Fault History pages of the CMCS. Ground tests can a.l so be performed u sing the CMCS. To access AFDS test functions, go to the Ground Tests page of the CMC menu (Figure 6-9-21), select chapter 22 AUTOPILOT FLT DIR and choose the appropriate test from the menu. Table 6-9-1 is a list of the tests available throug h the AFDS ground tests menu. Severa! of the tests have p reconditions th at must be me t befare the tests can take place. Preconditions are listed on the control display unit after the test selection has been made (Figure 6-9-21). CAUTION: Whenever performing operatíonal tests on any autopilot be sure the aircraft is clear of personnel ' and machinery.

1

6-47

TESTS AVAILABLE THROUGH THE AFDS GROUND TEST MENU l. L/ R/ C FCC

Tests the FCCs and the systems/ sensors that interface the the FCCs

2. MCP test

Tests the displays, switches, and control of the MCP

3. Ai leron servo

Command and engag e signal are sent to the aileron servo, the FCCs monitor the response

4 . Elevato r servo

The elevator servo is tested same as aileron servo test

5. Rudder servo

The rudd er servo is tested same as t he aileron servo test

Tests function of A/P disconnect 6. Autopi lot disconnect :witches switch 7. Go-around switches

Tests function of G/ A switches

8. Autoland unique test

Tests the operation of severa! functions critica! to th e autoland function

9. Air ground relay

Tests that all three FCC receive air/ ground data

1O. FCC configuration

Shows pin configuration of FCCs

11. FCC instrum ent

Monitors the interface between the FCCs and the integrated display system

12. Speed brake transducer

Monitors function and interface of S/B transducers

13. Flap transducer

Monitors function and interface of flap transducers

14. Stabilizer trim

The autopilot sends a given signal to the trim system and the FCCs monitor the response

15. Surface limit

Test to ensu re each FCC has t he same control surface travel limits (separate test conducted for aileron, rudder, elevator)

16. Tranducer output

Tests the stabi lizer, aileron, rudder, elevator, speed brake, and flap t ransducer outputs

Table 6-9-1. Autopilot flight d irector system (AFDS) ground tests

Many of the autopilot tests will opera te various control surfaces and/or thrust reversers. These control surfaces could cause damage to the aircraft or bodily injury to unsuspecting individuals. Also, be sure that other maintenance being performed on the aircraft will not adversely affect the autopilot tests and crea te a potential hazard. For example, if another technician is servicing the hydraulic system, the autopilot functional test should not be performed.

YawDamper The B-747-400 yaw damper system provides dampening for Dutch rol! prevention, turn coordination, and su ppression of structural

6-48

1

Autopilot and Autojlight Systems

D

•

2/8 GROU ND TESTS <21 ZONE TEMP CONTROL <22 AUTOPILOT FLT DIR

D

<22 YAW DAMPER

D D

<23 COMMUNICATIONS

D


<23AUDIO

1 D D D D

D D

D D D D D D

TEST PRECONDITIONS FCC-L

1/ 1

- SET THE L, C, ANO R I RS MODE SELECT SWITCHES ON PS TO ALIGN OR NAV.

--- -------------STARTTEST>


'

D D D D D D

,---

• •

AUTOPILOT FLT DIR 1/4

D D

'

D D D D D D


D D


TEST PRECONDITIONS MCP

D D D D D D

1/ 1

-NOTE: VISUAL VERIFICATION IS REQUIRED AS DESCRIBED IN THE SUBSEQUENT TEST PAGES.

- - - -- - ---------


STARTTEST >

D D D D D D

Figure 6-9-21. Sequence of AFDC ground test displays showing test preconditions Courtesy of Northwest Air lines, lnc.

~IR Us 1--------{~~o FCEPSHs

YAV U IIPU

1Éi11i11

/ il

r

CONTROl PANEl EIUs

AIR/GROUN D SYSTEH

RUOOER PEDAl INPUT

Figure 6 -9-22. Block diagram of the yaw damper system

modal .oscillations. Structural modal oscillations are an undesired e ffect created by turbulence, which cau ses bending of the fuselage around the w ing area. There are two redundant yaw damper systems, each containing a yaw damper module powered by the FCEPSM (fli ght control electronics power sup-

Courtesy of N orthwest Airlines, /nc.

ply modules). As seen in Figure 6-9-22, each yaw d amper module receives inputs fr om the IRUs, ADC, dedicated modal accelerometers, the yaw damper control panel, air/ground systems, the CMCS, hydraulic pressure sw itches, and a feedback signa! fro m the yaw damper actuators.

Autopilot and Autoflight Systems The main yaw damper modu le outputs go to the yaw damper actuators. Output signals also go to the EFIS/EICAS interface units (EIU) and the central maintenance computer system. The yaw damper control panel also receives an output from the yaw damper modules to verify the cu rren t operation of the yaw damper system.

Section 10

Fly-by-Wire The basic concepts of Fly-by-Wire (FBW) are simple; replace cables, pulleys, and pushrods with electrical w iring as a means to connect pilot inputs to aircraft control surfaces. In a traditional system, the pilot moves a control wheel, or yoke, and a stainless steel cable is used to transfer this motion into control surface movement. On most large aircraft, the pilot would move the control wheel, the cable would move a hydraulic actuator, and the hydraulic act uator would move the control surface. As seen in Figure 6-10-1 in a FBW system the pilot would move the control wheel, an electrical signa] would be sent to an electronics control unit comp uter, a nd the control unit would send an electrical signa! to the hydraulic actuator that moves the control surface. A FBW system mus t also employ a feedback system to provide a "realistic feel" back to the pilot through the flight deck controls (wheel, yoke, or rudder ped als). FBW is not actually a new concept. For years many aircraft ha ve employed electrica 1 circuits to operate certain control su rfaces. For example, Cessna ligh t aircraft, like the 172, have employed electric flap actu ators for severa! decades. The pilot would simply select a flap position using a switch on the instrument panel, the signa! would be sent to the flap motor and the flaps would move to the desired position. On sorne aircraft there was even a r udimentary feedback system, w hich would move an indicator to inform the pilot of flap position. The difference between this simple electric actuator and modern FBW systems is that today's aircraft use electrical signals to move primary flight controls, such as elevators and ailerons. Primary flight controls require con stant rep ositioning by the pilot and therefore require a m uch more complex system. A modern FBW design permits a more efficient a ircraft structure through the use of computer-aided controls. Th is technology allow s the airplane to meet strict safe ty requirements while decreasing weight and

increasing fue! efficiency. Modern FBW aircraft require a complex flight control system employing severa] computers, digital data transfer, and multiple act ua tors for dozens of fl ight controls.

B -777 Automatic Flight Control System The Boeing B-777, placed in service in the mid 1990s, was the first transport category aircraft designed to incorporate a fly-by-w ire primary flight control system. The FBW design had been employed on sorne militar y aircraft prior to the B-777 release; and is also used on the newer B-787 and A-380 aircraft for its reliability, efficiency, and weight savings. There are three distinct segments of the B-777 automatic flight control system: l. Flight management computing system (FMCS) 2. Autopilot flight director system (AFDS) 3. Flight controls and related mechanisms The B-777 flight controls are actually divided into two separate systems: the primary flight control system (PFCS) and the hig h lift control system (HLCS). Like most large aircraft, the B-777 employs dozens of independent flight controls (Figu re 6-10-2). As the name implies the PFCS is used as the primary control system providing both automatic and manual operations. The PFCS monitors a variety of inputs, employs various computers, and determines how and when to move control surfaces. The PFCS calculates commands to control surfaces using sensor inputs from control wheel, control col umn, rudder pedals, speed brake lever, and the pitch trim wheel. All three axes (pitch, roll, and yaw) a re provided stability augmentation and envelope protection by the PFCS. Envelope protection is used to ensure the airc¡·aft never exceeds the operational limits and enters into an unsafe configuration, such as a stall condition. The PFCS controls two ailerons, two flaperons, and fourteen spoilers for roll control; two elevators and a movable hori zontal stabilizer for pitch control; and a segmented (tabbed) rudder for yaw control. The HLCS is used to increase aircraft lift during takeoff and landing (low speed flight). The high Iift control surfaces include one inboard and one outboard trailing edge flap on each w ing. The B-777 also employs seven leading edge slats and one Krueger flap on each wi ng. A Krueger flap d iffers from an ordinar y flap or

1

6-49

6-50 1 Autopilot and Autoflight Systems ,

(A) BELL

CONTROL SURFACE

(RANK PUSH

Roo

o

(B)

HYDRAULIC ACTUATOR

DtRECT FORCE

Fig

me em

Th

MECHAN/CAL LINK

COl

anl

di.g A sig¡

(PC atel tor. (C)

SUii

PO

ElECTRICAL TRANSDUCER

ElECTRONIC CONTROL UNIT

HYDRAULIC ACTUATOR

DtRECT FORCE

FEEDBACK StGNAL

Figure 6-10-1. Operation of flight controls: (A) A mechanical flight control system, (B) A mechanicalhydraulic fli ght co ntrol system, (C) A fly-by-wire electronic-hydraulic flight control system

a slat in that the Krueger flap deploys from the leading edge of the wing and hinges from the front edge dow nward to increase lift. As seen in Figure 6-10-3, as the flight crew moves the wheel/yoke assembly, rudder pedals or other fligh t deck controls the movement is converted into an electrical signa! by position transducers. A transducer changes mechanica1 motion into an elect rical voltage. The electrical

sig na! is sent to the Actuator Control Electronic units (ACE). The ACEs convert the analog signa! from the transdu cers into a dig ital format and send that data to the pr imary flight computers (PFC) th rough an ARINC 629 data bus. There are three 629 buses dedicated to flight control data. The PFCs receive data from other systems, including the airplane information manage-

The whi fee<

PO WI. initi The the vate lot r tor í conf m an

Sino pm tion

Ase are r gene othe able

Autopilot and Autojlight Systems

1

6-51

ElEVATOR

ÜUTB0ARD FlAP

Figure 6-10-2. The fl ig ht controls of a B-777 aircraft

ment system, the air data and inertial reference units, and secondary air data sources. The PFCs con sider the input data and em ploys control-law software to calculate augmentation and envelope protections. The PFCs then send digital command signals back to the ACEs. The ACEs conver t the d igital sign als into analog signals for command of the power control units (PCU). Each PCU contains an electrically operated ser vo-valve that controls hydraulic actuators to move the contr ol surface. Each control surface will be connected to one, two, or three PCUs depending on load demands. The PCUs also contain a position transducer, which sends a feedback sign a! to the ACEs. The feedback signa! is used to determine when the PCU shou Id stop control surface movement. When in autopilot operations the AFDCs will initiate al! signals for control surface movement. The primary flight control system responds in the same manner as if the pilot m anually activated the flight deck controls. When in autopilot mode a signa! is sent to a backdrive actuator in arder to move the necessar y flight deck controls into the appropriate position as commanded by the autoflight computer. Since this aircraft is highly reliant on electrical power for flight control operations, interruption of electrical power cou ld be catastrophic. As discussed in Chapter 2 of this text, ther e are multiple engine-driven permanent magnet generators, a ram air backup generator, and other power sources including batteries available for flig ht control operations. Fig ure 6-10-4

FUGHT (OMPARTMENT

PRIMARY FUGHT

.:~-~-+'+!---~~+---++--++-- DECK (ONTROl S

MECHAN ICAl - -- - - - - - - - - - - (ONNECTION

~

--- ~ ELEVATOR, AILERON, RUDDER

Figure 6-10-3. The flight deck control movement is converted into an electrical signal by position transducers and converted into a digital signa! for use by the autoflight system.

6-52

1

Autopilot and Autoflight Systems

LEFT 28 V D C Bus 1

1

1

LEFT ENGINE GENERATOR (2 PMGs)

PSA- l

--

PFC-R

¡:

28 V DC CAPT FLT INST Bus

BATIERY

ACE-C ANO ASSOCIATED ACTUATORS PFC-C

H OT BATIERY Bus

P SA-C

-

-

¡.--

BATIERY '

ELECTRICAL CONTROL MODULES

-

r-

r-

AC E- l2 ANO ASSOCIATED ACTUATORS

r -- · 1 1

~--

:: 1

1 1

RIGHT 28 VDC Bus 1

ACE-R ANO ASSOCIATED ACTUATORS

•

:: ;

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ACE-l1 ANO ASSOCIATED ACTUATORS

PSA- R •

RIGHT ENG INE GENERATOR (2 P M Gs)

f-

r----

-

1 1 1 1

1 1 1 1 1 1

BATIERY

6-10-4. Redundant power sources provide safety for the B-777 flight control system.

shows the various sources available to power the ACEs and PFCs in the system. It is critical that these LRUs continue to receive power in order to maintain manual flight control. The primary flight control system can operate in three distinct modes: normal, secondary, and direct; all of which are dependent on the health of system sensors, computers, and control devices. The system w ill automatically sw itch to a less automated mode if cer tain components fai l and flight safety cannot be en sured. The normal mode of operation provides all envelope protections including stall w arning, over-speed, over-yaw, a nd bank angle. The au topilot is also fully functional in the normal mode. If there are one or more critica! failures the system soft ware will switch to secondary mode that limits sorne of the automated protections. If additional failures occur, the systems must operate in direct mode and only m anual pilot commands are accepted. The PFCs are non-operational in the direct mode. A big consideration for any fly-by-wire system is redundancy. For the most part, each flight deck control contains up to three pressure tran sducers to ensure pilot commands create the correct electrical signals. Of course, there are also multiple computers and redundant ~ software function s allowing the PFCS to fail fully active under most conditions. Of course if a failure is too extensive, the system will switch

operational modes from normal to secondary or direct. Fly-by-wire flight control systems also require a relatively complex feedback system to provide each pilot w ith the correct feel on the flight deck controls. This is important when pilots fly the aircraft manually. The B-77/ incorporates centering mecha nisms, which returns the control (wheel, yoke, and/ or rudder pedals) to the neutral position when appropriate. As seen in Fig ure 6-10-5 the system also incorporates electrical actuators, which return pressure as the pilot pulls/pushes or rotates a flight deck control. This return pressure will change with ,ürcraft speed and as the aircraft reaches flight envelope limits. Each of these feedback systems rely on a variety of inputs and severa! computer functions to provide pilots a "natural" feel when controlling the aircraft.

B-777 Autopilot Flight Director System The autopilot flight director system (AFDS) is an integral part of the B-777 flight controls. The AFDS has three ch annels that can each operate independently to provide redundancy. When activated, the autopilot function of the AFDS will control the aircraft on its selected vertical and horizontal flight path and selected airspeed. The flight di rector portian of the AFDS

6-10-5_ is opeG powerel mands • the p · the aira The AFl mode q flight c:fu ous seru provide 6-10-6 sb ponents.. pilot coc¡ cellanem to selea pilot lila] lateral na (VNAV' troJ swi¡¡¡ (TOGA; " and the co-pilot's

Autopilot and Autoflight Systems FUGHT (OMPARTMENT

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1 1

1 1

---------1 1

1

1

r----------· -----L---------·---------------1 ---~----.-----·---~----.----------1 1 1

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6-10-5. Two backdrive actuators position the fl ight de e k controls during autoflight operations. is operational whenever aircraft systems are powered that provides visual guidance commands on the aircraft flat panel displays giving the pilots al! the data needed to manually fly the aircraft. The AFDS has three major components: the mode control panel (MCP), three autopilot flight director computers (AFDC), and the various sensors, switches and transducers, which provide input signals to the system. Figure 6-10-6 shows the relationship of the AFDS comp onents. Here it can be seen that AFDS receives pilot commands through the MCP and the miscellaneous control switches. The MCP is used to select the operational mode of the AFDS. The pilot may select various operations, such as, lateral navigation (LNAV), vertical navigation (VNAV) and others. The miscellaneous control switches include: the takeoff, go-arou nd (TOGA) switches located on the throttle levers and the disconnect switches on the pilot's and co-pilot's control wheel.

The AFDS monitors the variou s !:"lilot activated inpu ts u sing three AFDCs. Each comp uter calculates the necessary response and sends output signals through th ree ARINC 629 data bus cables to the ACEs and PFCs. S.i milar to the manual flight operation s the ACEs and PFCs actívate the appropria te control surface. As a control surface is moved the PFC software calculates the backdrive commands, w hich are sent to the AFDCs. The AFDCs then send the backdrive signals to the appropriate backdrive actuators that reposition the rudder p ed a ls and /or control wheel/ yoke as needed. Most autopilot comm ands are redundant and the com puters an alyze multiple inputs. Software fu nctions known as mid-value selection and voting are used to d etermine the validity of all data prior to moving any control surface. To inform pilots of the current operating status, the autopilot flight director system will send data to the aircraft instrument display

1

6-53

6-54

1

Autopilot and Autojlight Systems FUGHT ( OMPARTMENT

~---------

~I RP. I!A NE

SENS0RS

1

1 1 1 1 1 1 1

_ _ __ _

MECHA NICAL _ _ __ _ _ _ _ ( ONNECTION

--- ~ ELEVATOR, A ILERON, RUDDER

AH AiJ S}'l

Figure 6-10 -6. Simplified diag ram of the B-777 autopilot flight director system (AFDS) system. The aircraft's PFDs, EICAS, and MFD will each show AFDS displays and a nnunciations as needed. The PFDs show flight modes as well as autoland and au topilot ind ications. All warning and caution data is sent to EICAS and the MFD shows AFDS status.

B-777 Flight Management

Computing System To help reduce pilot workload, a fl igh t management comp uting system (FMCS) is used to provide vertical and lateral gu idance for all ph ases of flight excluding takeoff and landing. The FMCS will also automatically tu ne all radios and provides navigation al data on th e flight deck d isplays. The FMCS software, known as the flight man agement computing function (FMCF) is located in the two AIMS cabinets. The B-777 airplane information management system was d iscussed in Chapter 3. One of the fl ight management computing fu nctions operates in active mode wh ile the other is ready in standby in the event of a failure. The flight management system operates

in conjunction w ith the autopilot flight d irector syste m to provide complete navigation and autofl ight functions. A simplified diagram of the FMCS is shown in Figure 6-10-7; please reference th is d iagram during the following discussions. The fligh t crew interface for the FMCS is through the three control d isplay units (CDU) located on the flight deck. The CDUs are mounted on the center pedestal between the two pilots and contain the trad itional alphanumeric keys and liqu id crystal display. The pilots enter all flight planning data on the CDUs and this infor mation is sent to both AIMS. The FMCF h as four basic elements: n avigation, flight planning, performan ce man agemen t, navigation radio tuning. Th e FMCF contains a large n avigational da tabase w ith all necessary navigational aids, wayp oints, flight pla ns, and other necessary information. Flight planning functions use flight crew inp uts to create the desired flight plan . The perfor mance ma nagement function employs

B-T

Co Thru! tion a Both

Autopilot and Autojlight Systems

1

6-55

, A/T Disconnect Switch es

Systems ARINC 629 Bus (4)

Throttles Intercabinet ARINC629 Bus

All

Airplane Systems

•:~ ;¡¡ • F/0 DSP

ISSP (2)

AIMS Cabinet (2)

Figure 6-10-7. A simplified diagram of the B-777 Flig ht Management Control System (FMCS)

aerodynamic models and flight crew selections to calculate the most economical flig ht path and engine power settings. The navigation radio tune function sets all radio frequencies and settings necessary for complete navigation for the entire flight. The FMCF software is updated at regular intervals in order to ensure currency.

B-777 Thrust Management Computing System Thrust m anagement is an independent function also contained in the two AIMS cabinets. Both AIMS contain redundant system s to ensure fail active operations. The thrust management computing function (TMCF) is basically software used to send auto throttle commands to the throttle servo motors and engine trim commands to the engine electronic controllers (EEC). The servo mo tors are used for large throttle adjustments and the EECs make fine adjustments to keep both engines at peak efficiency for various flight conditions.

A-380 Flight Control System The Airbus A-380 is a large four-engine transport category aircraft employing state of the art integrated electronics and the AFDX data transfer system. AFDX was discussed in Chapter 2. The aircraft uses an advanced flight control system called the Auto Flight System (AFS). The A-380 AFS can be divided into three distinct elements: flight g uidance (FG), flight management system (FMS), and the flight controls. Due to the size and complexity of the A-380, the fligh t control system contains nearly 50 separate control surfaces activated automatically or m anually by pilot commands (Figure 6-10-8). The A-380 employs a fly-by-wir e-type system with all flight deck inputs converted to electrical signals, rou ted through one or more processor circuits and eventually sent to an electrically controlled hydraulic actuator assembly (Figure 6-10-9). The flight controls are div ided into two distinct catcgories: pri-

6-56

1

Autopilot and Autojlight Systems

HORIZONTAL STABILIZERS - -

c=J

. . LIFT DEVICES

PRIMARY FUGHT (ONTROLS

Figure 6-10-8. An A-380 flight control system co ntains nearly 50 separate co ntrol surfaces.

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-

SPEED BRAKE

PRIM

(3)

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-

FWS

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RUDDER PEDALS

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re

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Figure 6-1 0-9. A sim plified d iagram of the A380 fly-by-wire system

If a all anQ

mary flight controls and the slats and flaps. The primary flight controls are used for control of pitch, r oll, and yaw during normal, direct, or alternate flight configurations. The slats and flaps are each considered a high-lift device u sed for low speed flight d uring takeoff and landing. The primary flight control system employs three primary com puters (PRIM), which provide flight control, flight guidance, and envelop pro-

tection functions. Envelope protection is provided by system software to prevent exceedance of certain flight parameters, such as, excessive bank angle. The system also employs three secondary computers (SEC). Each computer, PRIM and SEC, can perform two functions: command computations and command executions.

~~

Command computations convert pilot or autopilot commands into contr ol surface deflection signals according to flight parameters and

a no are l fligfj

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the dm whe

Autopilot and Autoflight Systems

r----i----, PRIM 3

r1

__ i _ SEC 1

SEC 2

Figure 6-10-1O. The A-380 flight control system operating in normal m o de

The flight control system operates in three d istinct modes: normal, direct and alternate. In norm al m ode one primary computer perfor ms all computation functio ns and sends command signals to the o ther computers (Figure 6-lC-10). All three PRIM a nd SEC computers perform the execution functions for their assigned control s urfaces. As the d iagram shows, the PRIM m aster computer also p er for m s self-monitori ng of the control surface feedback s ig nal to ensure the system s respond acc urately. If a malfunction is detected in the master PRIM all computation functions w ill be passed to another PRIM. If all PRIMs are lost d ue to failures, each SEC w ill perform computation and execution functions as needed. At this point the flig ht control system w ill automatically downgrade to direct mode. Direct mode occurs whenever the system has degraded d ramatically due to severa! failures and the normal mode of operation is not available. When operating in direct mode the auto trim function is no longer available and all envelope protections are lost. Warning inform ation displayed on the flight deck such as, over speed or stall warn-

6-57

r ____ ( ____• r-

AIRCRAFT FEEDBACK

envelop protection limitations. In addition, command computations analyze and compare servo actuators feedback signals in order to ensure proper control surface movement h as been achieved. The command execution function of PRIM and SEC computers send the necessary electrical control signals to the servo actuators in order to create control surface movement.

1

ings, inform the pilot of any potential envelope exceedance. If all PRIM a nd SEC computers are lost due to system failures, the aircraft flight controls are operated in alternate or backup mode. The backup system is totally segr egated from the normal system with dedicated sen sor s and transducers in the flight deck controls. At least one hydraulic system and backup electrical power source must be available for alternate mode operations. If all engines fail power is provided by the ram air turbine. In alternate mode only flig ht controls used for basic maneuvers and safe landing are available. Figure 6-1011 shows the control surfaces that are operable in the direct mode of operation. The A-380 flight control system employs three ty pes of servo actuators to move the flight control surfaces and high lift devices. The actuators are combinations of electron ic controllers, electric motors, and hydraulic actuators. Since this is a fly-by-wire aircraft, each servo is electrically controlled from one or more computers. There a re three types of servo actuators: conventional actuators, electro-hydrostatic actuators (EHA), and electrical backup hydraul ic actuators (EBHA) .. As seen in Figure 6-10-12 conventional servos employ an electrically controlled ser vo valve that regulates the flow of hydraulic fluid into the actuator and determines control surface movement. The servo · val ve can also be u sed to reverse the flow of hydraulic fluid, w hich ch anges the actuators direction of travel. In

__i___ _ SEC 3

6-58 1 Autopilot and Autoflight Systerns

IN BOARD AILERON

/

HORIZONTAL STABILIZERSINBOARD ElEVATORS

/

INBOARD Al LERON

/

Figure 6-10-11 . A-380 flight controls operated in direct mode order to operate, conventional serves must have a supply of pressurized hydraul ic fluid. The A-380 employs multiple centralized hydraulic systems to provide redundancy for the actuators. Each system is named by color, green or yellow. The electro-hydrostatic actuators are hydraulic units that have their own electric motor and self-contained hydraulic system. The servo receives an electronic signa! to its electric motor located within the actuator assembly. The electr ic motor is directly coupled to a hydraulic pump that supplies the pressurized fluid to move the actu ator. This type of actuator is independent of the central hydraulic systems, but requires a supply of electrical current to drive the pump motor. The electrical backup hydraulic actuator is a combination of the conventional and electrohydrostatic actuators. This unit is connected to the central hyd rau lic system using a servo valve for electronic control a nd employs a selfcontained electric motor/pump assembly to produce an independent supply of hydraulic pressure. The backup actuators can therefore operate using electric or hydrau lic power.

ance by sending targets, such as waypoints, ai rports, and navigation aid s, to the FG system. The AFS works in conjunction with the three PRIM computers for autopilot, flight director, and auto throttle functions. There are two complete FMS functions, which operate using one of three FMCs, providing redundancy to ensure that systems can fail and the FMS still remains operational. A basic auto flight system interface diagram is shown in Figure 6-10-13. The flight crew can interface w ith the AFS u sing the AFS control panel or the three MFDs can provide back up for the control panel. The MFDs operate in conjunction w ith the KCCU (keyboard and cursor control unit) as discussed in Chapter 3. The PFDs provide visual feedback to the pilots regarding the AFS operations. The NDs show all n avigation data related to AFS. Of course there are discrete controls such as autopilot disconnect switches on the side stick control and auto throttle disconnect switches on the throttle quadrant.

Fig an

The The FG function of the AFS is designed to follow short-term instructions and provide guidance and speed controls. The FG functions include: l. Autopilots one and two (APl and AP2) pro-

vide calculations for pitch, roll, and yaw

A-380 Auto Flight System The Airbus A-380 Auto Flight System (AFS) is comprised of two distinct elements; the flight guidance (FG) and the flight management system (FMS). The FG system provides shortterm lateral and vertical guidance based on the flight parameters selected by the flight crew or the FMS. The FMS provides long-term guid-

2. Flight directors one and two (FDl and FD2) provide guidance commands on the PFDs. This en ables the fligh t crew to manu ally fly the aircraft or to monitor guidance orders during au topilot controls 3. Auto thrust (A/ THR) controls engine thrust through the FADEC (full authority

~ :~d FMq 13).

the

r~

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Autopilot and Autojlight Systems

( ONVENTIONAL A CTUATOR

1

6-59

ELECTRO - H YDROSTATIC A CTUATOR

( 8ACKUP SYSTEM)

A380 FUGHT ( ONTRO LS

ELECTRI CAL 8 ACKUP H YDRAULIC A CTUATOR

Figure 6-10-12. The A-380 employs three types of servo actuators, conventional, electro-hydrostatic and electrical backup hydraul ic digital engine control). The FADEC system analyzes the thrust command and operates electrical servo actuators that change engine power settings and position the engine throttles accordingly The pilot interface to the AFS is through the traditional flight deck controls and instrument display system found on the A-380. It is important that fly-by-wire aircraft retain a "traditional feel" so a pilot can easily transition from one aircraft to another. The th ree main inputs to the AFS are the autoflight system control panel (AFS CP), the MFD and KCCU, the thrust levers, and the side stick control. Each of these input devices creates an electric signa!, which connects to the PRIM, FMC, or FCU back up computers (Figure 6-1013). The computers also send feedback signals to the flight deck controls which are employed to provide feedback to the pilot. The PFD, D, and MFD provide indications as to the AFS system status.

• PFD

B

ND

l 1

-·

MFD

•

_!

KCC U

1

. SI DESTICK PUSH BUTTO N

¡----+

-

-

P RI M

+

-

FMC

THRUST LEVERS

1 1

- - ---FCU - -BACKUP -- - --- -

Figure 6-10-13. A basic A-380 autoflight system interface diagram