AUTOPILOT FLIGHT DIRECTOR SYSTEM

CH 22

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ATA 22 AUTOPILOT FLIGHT DIRECTOR SYSTEM AFDS GENERAL - COMPONENT LOCATION .....................................84
MODE CONTROL PANEL - FUNCTIONS ............................................86
AUTOPILOT FLIGHT DIRECTOR SYSTEM - INTRODUCTION ............4 AFDS - ANN AND WARN - GEN DESCRIPTION .................................88
TMS - GENERAL DESCRIPTION ...........................................................8 AFDS GEN - AUTOLAND STATUS ANNUNCIATORS .........................90
MAINT MONITOR SYSTEM - MCDP AUTO PWR ON MODE..............10 AFDS - FUNC DESC - ASA - DISPLAY SEQUENCE -1 ........................92
FLT CONT/HYD - GENERAL DESCRIPTION ......................................12 AFDS - FUNC DESC - ASA - DISPLAY SEQUENCE - 2 ......................94
FLT CONT/HYD - FLIGHT CONTROLS - INTRO .................................14 AFDC - FUNC DESC - ASA - DISPLAY SEQUENCE - 3 ......................96
COMPONENT LOCATION - FLIGHT CMPTMNT .................................16 AFDS - FUNC DESC - ASA - DISPLAY SEQUENCE - 4 ......................98
FLT CONT/HYD - COMPONENT - HYD SYS CONTROLS ..................18 MODE CONTROL PANEL INTERNALS .............................................100
FLT CONT/HYD - COMPONENT - FLT CNTRL INDICATORS ............20 ENGAGE LOGIC - SIMPLIFIED .........................................................102
FLT CONT/HYD - AILERON CONTROL SYSTEM ...............................22 ENGAGE LOGIC DETAILS - HARDWARE MONITORS .....................104
FLT CONT/HYD - ELEVATOR CONTROL SYSTEM ............................24 INTERFACE - FCC CROSS-CHANNEL DATA ...................................106
FLT CONT/HYD - RUDDER CONTROL SYSTEM ................................26 AFDS POWER DISTRIBUTION ..........................................................108
YSM - YAW DAMPER INTRODUCTION ...............................................28 AFDS FCC POWER DISTRIBUTION .................................................110
YSM - YAW DAMPER - GENERAL DESCRIPTION .............................30 AUTOLAND POWER SWITCHING - FUNC DESC ............................112
YSM - YAW DAMPER - COMPONENT LOCATION .............................32 AFDS GENERAL - POWER ISOLATION LOGIC ................................116
YSM - GENERAL DESCRIPTION .........................................................34 AUTOLAND SEQUENCE ...................................................................118
YSM - GENERAL DESCRIPTION - FUNCTIONS .................................36 THRUST MANAGEMENT SYSTEM - INTRODUCTION ....................120
YSM - FAULT RECORDING ................................................................38 TMS - COMPONENT LOCATION .......................................................122
YSM - BITE ...........................................................................................40 TMS - THROTTLE COMPONENT LOCATION ...................................124
YSM - FAULT ANNUNCIATION ............................................................42 TMS - THRUST MODE SELECT PANEL ...........................................126
YSM - YAW DAMPER - PRE FLT TEST FROM FLT DECK .................44 TMS - OP- EADI DISPLAYS ...............................................................128
YSM - STAB TRIM - INTRODUCTION ..................................................46 TMS - OP- ENGINE EICAS DISPLAY - THRUST LIMIT FUNC ..........130
YSM - STAB TRIM - COMPONENT LOC - FLT DECK .........................48 TMS - FUNC DESC - SYSTEM BLOCKDIAGRAM ............................132
YSM - STAB TRIM - CMPNT LOC - MEC AND JKSCRW AREA ..........50 TMS - FUNCTIONAL DESC - SYS BLOCK DIAGRAM CONT. ..........134
YSM - STAB TRIM - GENERAL DESCRIPTION ..................................52 MAINT FUNCTIONS ...........................................................................136
YSM - STAB TRIM - FUNCTIONAL MODE PRIORITY .........................54 MAINT MONITOR SYSTEM - REMOTE MCDP OPERATION ...........138
STAB TRIM SYS - MNL ELEC STAB TRIM - DESCRIPTION .............56 MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS .......140
STAB TRIM SYS - ALT ELEC STAB TRIM - FCTNL DESC .................58 MAINTENANCE CONTROL AND DISPLAY PANEL ..........................144
YSM - STAB TRIM - FCTNL DESC - AUTO STAB TRIM... ...................60 MAINTENANCE MONITOR SYSTEM - COMP LOC ..........................148
YSM - STAB TRIM- FAULT ANN - STAB TRIM ....................................62 REMOTE MCDP CONTROL PANEL CONNECTOR LOC ................150
YSM - STAB TRIM - FAULT ANN - UNSCHEDULED TRIM .................70 LRU FAULT CONSOLIDATION ..........................................................152
FLT CONT/HYD - GEN DESC - FLIGHTCONTROLS ...........................72 FAULT MESSAGE FORMATS ...........................................................154
AUTOPILOT CONTROL SERVO ..........................................................74
LATERAL CENTRAL CONTROL ACTUATORS .................................76
AFDS GENERAL - AFDS - GENERAL DESCRIPTION .......................78
AUTOPILOT FLIGHT DIRECTOR SYSTEM .........................................80
AFDS GENERAL - BLOCK DIAGRAM ..................................................82
 STUDENT NOTES:

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AUTOPILOT FLIGHT DIRECTOR SYSTEM - INTRODUCTION

Thrust Management System

The thrust management system has two functions. The system moves the
thrust levers and calculates the thrust limit for the EICAS display.

Yaw Damper System

The yaw damper system controls the rudder to decrease yaw oscillations
because of a Dutch roll or gustinduced sideslips.

Automatic Stabilizer Trim and Mach Trim System

The automatic stabilizer trim and Mach trim system controls stabilizer position
as a function of these:

- Mach at high speeds

- Autopilot trim commands with the autopilot engaged.

Autopilot Flight Director System

The autopilot flight director system gives automatic control for these control
systems to operate the selected mode:

- Aileron
- Elevator
- Stabilizer
- Rudder.

The autopilot flight director system also gives pitch and roll flight director
commands, system warnings, and annunciations.

Maintenance Monitor System

The maintenance monitor system gives a centralized flight and ground faults
readout and ground test function for these systems:

- Autopilot flight director system

- Thrust management system
- Flight management system.
 AUTOFLIGHT SYSTEMS

THRUST YAW DAMPER AUTOPILOT/FLIGHT AUTO STAB TRIM

MANAGEMENT SYSTEM DIRECTOR SYSTEM SYSTEM
SYSTEM (DUAL) (TRIPLE) (DUAL)
(SINGLE)

MANUAL MANUAL MANUAL MANUAL

INPUTS FMCS INPUTS INPUTS INPUTS

MAINTENANCE ROLL CONTROL - FLIGHT DIRECTOR PITCH CONTROL

THRUST YAW
MONITOR SYSTEM COMMANDS SYSTEM
CONTROL CONTROL
SYSTEM (AILERONS, - AFDS ANNUNCIATIONS (STABILIZER
(THROTTLES) (RUDDER)
(MCDP) SPOILERS) AND WARNINGS AND ELEVATOR)

AUTOPILOT FLIGHT DIRECTOR SYSTEM - INTRODUCTION

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AUTOPILOT FLIGHT DIRECTOR SYSTEM - INTRODUCTION

General

The autopilot flight director system (AFDS) is a triple-channel system which

operates as a single channel system for cruise. For cruise, the AFDS controls
the pitch and roll axes of the airplane and gives pitch and roll flight director
commands. For approach, land, rollout, and go-around, the AFDS operates as a
multi-channel (dual or triple) system to control the pitch, roll, and yaw axes of
the airplane.
 AUTOPILOT FLIGHT DIRECTOR SYSTEM - INTRODUCTION
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TMS - GENERAL DESCRIPTION

Purpose

The thrust management computer (TMC) has a thrust limit function and an
autothrottle function. These functions each have different modes that you select
and show in different locations.

Thrust Limit Functions

Selection of the thrust limit modes are from the thrust mode select panel
(TMSP) or by the FMC in VNAV mode. They show on EICAS.

The thrust limit functions are always active. The thrust limit shows on EICAS
and is an upper limit for autothrottle calculations.

Autothrottle Functions

Selection of the autothrottle modes are from the AFCS mode control
panel(MCP) or by the FMC in VNAV mode. They show on the EADI.

The autothrottle engages in a mode when the AFCS MCP switch is in the A/T
ARM position, and you select a mode. These functions move the thrust levers.

The propulsion interface management unit (PIMU) is a digital buffer between

the electronic engine control (EEC) and the TMC.
 TO
CLB 1 2
GA

TEMP SEL TAT + 12c

CON CRZ
THRUST CRZ
LIMIT
FUNCTIONS
THRUST MODE
SELECT PNL

FMCS UPPER EICAS DISPLAY

SPD
A/T ARM
IAS/MACH
F/D L NAV
IAS
ON OFF

SEL
THR V NAV
OFF

SPD FL CH
AUTOTHROTTLE
FUNCTIONS
AFCS MODE CONTROL PANEL ADI
INTERFACING SYSTEMS
GO-AROUND SWITCHES
AUTOTHROTTLE DISENGAGE SWITCHES

THROTTLE LEVER SERVO DRIVE

EECS PIMU
ANGLES (TLA) TACH FEEDBACK
AUTOTHROTTLE
THRUST MANAGEMENT SERVO
COMPUTER

TMS - GENERAL DESCRIPTION

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MAINTENANCE MONITOR SYSTEM - MCDP AUTO POWER Faults Not Stored

ON MODE Self-tests are DONE during the auto on mode. Any selftest failure will cause the
MCDP to shut down. Failure data for this condition is not stored. Failures will
Initiation
show when you do a manual power-up.
The MCDP comes on when either air/ground input changes from an in-air to an
on-ground state. This normally occurs at touchdown in order to store last flight
faults.

The STORE DATA Bit is sent by the computers only when this sequence of
events has occurred:

- IRS ground speed 100 Kts and air/ground = in air

- IRS ground speed 40 Kts and air ground = on ground
- For the FCCs, the autopilot channel is disengaged.

Termination

The auto power on mode will stop when last flight fault data has been received
from all computers with their store data bits set on, or after three minutes.
During the auto power ON mode, the MCDP switches are deactivated.

Undesired Power-Up

Auto power-up will also occur if the air-ground circuit breakers, LDG GR
POSITION AIR/GND SYS 1 and POSITION AIR/GND SYS 2, are cycled from
off to on because this sends a transition from in-air to on-ground to the
MCDP. Air/ground inputs may also cycle during power transfer of external/
generator/APU power. You may stop the three-minute auto power-up period, if
caused by ground operation of the air/ground circuit breakers or power transfer
by a cycle of the MAINT CONT DISPLAY circuit breaker.

Message Displayed

During this auto power-on mode, the display message is AUTO ON MODE IN
PROGRESS.
 FLIGHT
FAULTS
L FCC

C FCC

R FCC
AUTO ON MODE
IN PROGRESS
- 3 MINUTES
TMC OR
- ALL FLIGHT
FAULTS STORED
..... .. L FMC
MCDP
- AUTOMATIC
SHUTDOWN
AIR/GROUND RELAYS
SENSE TOUCHDOWN
R FMC

MCDP
- AUTO POWER ON
- STORE FLIGHT
FAULTS IF
STORE DATA
BITS SET

MAINTENANCE MONITOR SYSTEM - MCDP AUTO POWER ON MODE

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FLT CONT/HYD - GENERAL DESCRIPTION CAUTION: ACCIDENTAL DEPLOYMENT OF THE RAT MAY OCCUR WHILE
DOING GROUND TESTS OF THE PITOT SYSTEM.
Only one actuator gets power supply from only the left hydraulic system. DEACTIVATE THE SYSTEM BY PULLING THE RAT AUTO
CIRCUIT BREAKER ON THE P11 PANEL.

Continuous Operation
System Pumps Demand Pumps Conditions
Basic System
L or R EDP Pressure
System
L or R ACMP Pressure
During Large
Loads
Basic System
Center ACMP 1 Pressure
Basic System
Center ACMP 2 Pressure
(Auto Load
Shed)
System
Center ADP Pressure
During Large
Loads
Operates
Center RAT When
Extended
Hydraulic Power Summary

Training Information Point

Note: Each hydraulic system uses a fluid-to-fuel heat exchanger in the

main fuel tanks. There must be at least 600 gallons of fuel in the
left tank for operation of the left or center hydraulic systems and
600 gallons in the right tank for operation of the right system.
 LEFT HYD SYS - FLT CONT CENTER HYD SYS - FLT CONT RIGHT HYD SYS - FLT CONT

EDP AC AC AC AC
MP ADP EDP MP
MP MP

L E SLATS
RAT
T E FLAPS

STAB TRIM STAB TRIM

L WING SOV L TAIL SOV C WING SOV C TAIL SOV R WING SOV R TAIL SOV

LCCA LCCA LCCA

SPOILER PCA SPOILER PCA SPOILER PCA
(1-6-12) (3-4-5-8-9-10) (2-7-11)
AIL PCA AIL PCA AIL PCA
(LOB-LIB-ROB) 1 (LIB-RIB) 1 (LOB-RIB-ROB) 1
ELEV A/P ELEV A/P ELEV A/P
SERVO SERVO SERVO
ELEV FEEL ELEV FEEL
CMPTR CMPTR
ELEV PCA ELEV PCA ELEV PCA
(LEFT-RIGHT) (LEFT-RIGHT) (LEFT-RIGHT)
DIR A/P DIR A/P DIR A/P
SERVO SERVO SERVO

RUDDER PCA RUDDER PCA RUDDER PCA

RIGHT YAW LEFT YAW
DAMPER SERVO DAMPER SERVO 1 - LOB - LEFT OUTBOARD
- LIB - LEFT INBOARD
RUDDER RATIO - RIB - RIGHT INBOARD
CHANGER - ROB - RIGHT OUTBOARD

FLT CONT/HYD - GENERAL DESCRIPTION

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FLT CONT/HYD - FLIGHT CONTROLS - INTRODUCTION

Purpose

The flight control system controls the airplane attitude as necessary to follow a
flight profile.

Flight Control System Description

The flight control system includes primary controls which directly control
airplane attitude and secondary controls which change the effectiveness of
primary controls.

These are the primary controls:

- Two inboard and two outboard aileron surfaces provide roll or lateral control
- Two elevator surfaces provide primary pitch control
- One rudder surface provides yaw control.

These are the secondary flight controls:

- Twelve spoiler segments help roll control and effect lift and drag
- The moveable horizontal stabilizer helps pitch control
- Twelve leading edge slats and four trailing edge flaps are high lift devices
which change the effectiveness of the primary control surfaces.
 FLT CONT/HYD - FLIGHT CONTROLS - INTRODUCTION
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COMPONENT LOCATION - FLIGHT COMPARTMENT

Primary Flight Control Inputs

Manual inputs for pitch, roll, and yaw control are provided by the captain and
first officer control column, control wheel, and rudder pedals. The control wheels
have switches to disengage the autopilot.

Aileron trim control is provided by two switches on the P8 aft aisle stand panel.
The trim position indicators are on top of the control wheels.

The rudder trim switch is a round knob to the right of the aileron trim switches.
The rudder trim position indicator is forward of the trim switch

Secondary Flight Control Inputs

Stabilizer trim control can be done either electrically with the thumb switches on
the outboard horn of each control wheel or with either alternate electric

stabilizer trim arm and control switches or via manual arm and control levers on
the P10 quadrant stand. The stabilizer trim is the only method for pitch trim.

There is no elevator trim. The manual flap / slat control lever and speed brake
control lever are also on the P10 quadrant stand.
 ALTERNATE SPEED BRAKE FLAP/SLAT
ELECTRIC TRIM CONTROL CONTROL
SWITCHES LEVER LEVER

STABILIZER TRIM STABILIZER TRIM

SWITCHES SWITCHES
A/P DISENGAGE
A/P DISENGAGE SWITCH
SWITCH
AILERON TRIM
AILERON TRIM POSITION
INDICATOR INDICATOR

P10 QUADRANT
STAND

STABILIZER
TRIM POSITION
INDICATORS

RUDDER TRIM
POSITION
INDICATOR

AILERON TRIM
CAPT SWITCHES F/O
RUDDER RUDDER
PEDALS RUDDER TRIM PEDALS
SWITCH
CAPT CONTROL COLUMN P8 AFT AISLE STAND PANEL F/O CONTROL COLUMN

COMPONENT LOCATION - FLIGHT COMPARTMENT

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FLT CONT/HYD - COMPONENT - HYDRAULIC SYSTEM

CONTROLS
Engine Ignition and Start Control Panel

The switch for manual deployment of the ram air turbine (RAT) is on this panel.
You push the switch to extend the RAT. The lower part of the switch
(UNLOCKD) comes on amber to show that the extend mechanism releases.
Electric power connects to a motor to extend the turbine. The upper part of the
switch (PRESS) comes on in green when the RAT extends, and the pressure
gets to 1275 psi.

You cannot retract the RAT in the air. You must manually put the blades to
center for retraction on the ground.

Stabilizer Trim Hydraulics Power Cutout Switches

The stabilizer trim hydraulic power cutout switches individually control left and
center hydraulic pressure to the stabilizer trim control modules. The switches
are on the P10 control stand.

Hydraulic Flight Control Panel

The hydraulic flight control panel has controls that isolate hydraulic systems for
troubleshooting. Push-on/push-off switches on the hydraulic flight
control panel individually operate hydraulic system shutoff valves (SOVs).
Shutters in the switch are a cover for or show the word ON in the upper part of
the switch. The word OFF comes on in amber when the SOV is off.

Wing SOVs control pressure to the wing actuators and servos for ailerons and
spoilers.

Tail SOVs control pressure to elevator and rudder components.

 FLT CONT/HYD - COMPONENT - HYDRAULIC SYSTEM CONTROLS
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FLT CONT/HYD - COMPONENT - FLIGHT CONTROL

INDICATORS
Trim Indicators

The aileron trim indicator is on the control column. The pointer is in the middle of
the control wheel. When you use the trim, the neutral position changes, and the
control wheel gets a backdrive to the new position. Units of trim show at the
pointer.

The stabilizer trim position indicator has an electrically driven pointer that
operates in response to a position transmitter. The position transmitter
moves by a cable attached to the stabilizer. The scale has colored segments to
show the takeoff range. The rudder trim indicator is almost the same as the
stabilizer indicator. The transmitter moves by mechanical linkage with the
rudder surface.

Trim Switches

The three-position aileron switches are spring-loaded to the center position and
momentary in LEFT WING DOWN and RIGHT WING DOWN positions. You
must use the two switches to give arm and control signals to the trim
actuator.

The rudder trim switch is spring-loaded to center and turns in the necessary
direction of trim.

Flight Control Surface Position Indicators

The position of each flight control surface shows on the EICAS lower display.
You push the STATUS switch on the EICAS control panel to see the display. All
status pages include the control surface position display.
 APL
NOSE 0
DN
2
S
AILERON TRIM T 4
A 6
6 4 2 0 2 4 6 B
T 8
R 10
I RUD
M O
12
APL F
NOSE
14
F
UP

AILERON TRIM
AIL ELEV AIL
POSITION INDICATOR
(2) (CONTROL
( WHEEL)
STABILIZER TRIM EICAS DISPLAY
POSITION INDICATORS
RUDDER TRIM (2) (P10)
POSITION
INDICATOR 15 10
RUDDER TRIM
5 0 5 10 15

NOSE LEFT UNITS NOSE RIGHT

AILERON TRIM AILERON

SWITCHES

LEFT RIGHT NOSE NOSE DISPLAY COMPUTER BRT THRUST REF SET
WING WING LEFT RIGHT ENGINE STATUS AUTO
BRT
BAL BOTH
EVENT L R R MAX IND
DOWN DOWN RECORD
L
RESET
RUDDER TRIM
SWITCH
AILERON/RUDDER TRIM
CONTROL PANEL (P8) EICAS CONTROL PANEL (P9)

FLT CONT/HYD - COMPONENT - FLIGHT CONTROL INDICATORS

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FLT CONT/HYD - AILERON CONTROL SYSTEM

Primary Aileron Control Functions

The primary control inputs are through the captain control wheel and lateral
central control actuators (LCCAs) to the aileron power control actuators (PCAs).
A feel, centering, and trim assembly conditions manual control inputs.

Back-up Aileron Control Functions

The back-up control inputs are through the first officer control wheel by forward
and aft bus cables to the primary system. Final back-up input can be direct
mechanical movement of the wing cable system through override and lost
motion devices.

Autopilot Control Functions

The flight control computers (FCC) receives inputs from the mode control panel
and other sources. The control surface commands go to the LCCAs which
provide mechanical movement to aileron PCAs.

Automatic Control Functions

The outboard aileron lockout device eliminates commands to the outboard

ailerons at cruise speeds. The inboard aileron droop device responds to trailing
edge flap movement to maintain control effectiveness and reduce drag by filling
the gap between flap segments.
 CAPTAIN FIRST OFFICER
CONTROL WHEEL CONTROL WHEEL

TO
RIGHT
WING

R LCCA OUTPUT
QUADRANT

CENTER LCCA

L LCCA

L LCCA OUTPUT R TORQUE

QUADRANT TUBE

FEEL CENTER
AND TRIM
L TORQUE ASSY
TUBE
R LCCA
TO
LEFT
WING
AILERON PCA
(8)

FLT CONT/HYD - AILERON CONTROL SYSTEM

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FLT CONT/HYD - ELEVATOR CONTROL SYSTEM

Control Inputs

Manual pitch attitude control inputs can be from either control column by a
separate cable system to the aft quadrant aft of the horizontal stabilizer. The two
are interconnected by a torque tube at the control columns and by linkage at the
aft quadrant. There is no power boost for manual inputs.

Autopilot input is from three flight control computers (FCC) to three servos. The
servos move the aft quadrant torque tubes. The servos have LVDTs which
provide autopilot actuator position and output position to the FCCs.

Control Conditioning

Manual inputs are conditioned by the feel unit which obtains data from the feel
computers. Feel pressure changes with airspeed and stabilizer position. The
feel unit also has a centering mechanism to maintain a neutral position when
there is no input.

Control Outputs

Elevator control movements go to the left and right power control actuators
(PCA) by mechanical linkage. There is an interconnection between elevator
PCA linkages by slave cable to prevent a large asymmetry. Position transmitters
are at each elevator to provide control surface position on the EICAS display.
 CONTROL
CAPTAIN COLUMN
OVERRIDE FIRST OFFICER OVERRIDE
CONTROL CONTROL COLUMN MECHANISM
COLUMN MECHANISM FEEL AND
CENTERING UNIT AUTOPILOT PITCH
TENSION CONTROL SERVO
REGULATOR (3) AFT QUADRANT
QUADRANT ELEVATOR INTERCONNECT
FEEL HORIZONTAL ROD
(2) STABILIZER
COMPUTER
SLAVE
STICK CABLE
SHAKER INTER-
(2) CONNECT

SLAVE CABLE
QUADRANT CENTER LINE OF
(2) STABILIZER REAR
SPAR HINGES
POSITION
TRANSMITTER
AFT QUADRANTS (2)
LEFT AFT QUADRANT OVERRIDE
OUTPUT ARM MECHANISM RIGHT AFT QUADRANT
CONTROL ROD OUTPUT ARM
CONTROL ROD
LEFT AFT QUADRANT RIGHT AFT QUADRANT
POWER CONTROL ACTUATORS (PCA(S) TORQUE TUBE TORQUE TUBE
(3 PLACES ON EACH
LOST MOTION AND OUTBOARD ELEVATOR) LEFT INBOARD ELEVATOR
OVERRIDE DEVICE (RIGHT SIDE SIMILAR)
(2) FWD
LEFT OUTBOARD ELEVATOR
(RIGHT SIDE SIMILAR)
OUTBD

FLT CONT/HYD - ELEVATOR CONTROL SYSTEM

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FLT CONT/HYD - RUDDER CONTROL SYSTEM

Trim inputs from the flight compartment control switch move the aft quadrant
Rudder System Description assembly and give a maximum rudder movement of 16.8 degrees. Trim
operation does a backdrive of the cables and rudder pedals.
Rudder system components are in different locations. Rudder pedals, the trim
switch, and indicator are in the flight compartment. System electronics are in the Power Control Actuators (PCA)
main equipment center. These components are in the vertical stabilizer:
Three PCAs move the rudder each with a different hydraulic system. The left
- Autopilot rollout guidance servos hydraulic system pressure to the middle PCA goes through the ratio changer
- Aft quadrant with the trim feel and centering unit actuator. If the ratio changer function has a failure, the middle PCA
- Ratio changer mechanism depressurizes. Each PCA has an override in the input linkage to its control
- Yaw damper actuators valve.
- Power control actuators (PCAs).

The position transmitter on the rudder gives position data for the EICAS display.

Rudder System Function

Manual rudder inputs are mechanical to the PCA control lever. These
components have an effect on the input:

- Feel, centering, and trim mechanism

- Ratio changer
- Yaw damper.

Autopilot inputs from the flight control computers give directional guidance only
during a multichannel approach and rollout.

The YSMs use airspeed and yaw rate from the ADIRU to give turn coordination
and decrease unwanted yaw. The rudder ratio module controls the ratio of
rudder input to rudder movement through the ratio changer. The schedule of
rudder ratio as a function of airspeed is from the yaw damper/stabilizer trim
module (YSM). All inputs move the PCA control levers to control rudder
movement. Each PCA gets power from a different hydraulic system.

Trim
 PCA (3)

(C)

YAW DAMPER MODULES YAW DAMPER

L AND R SUMMING
MECHANISM (L)

TEMPERATURE
COMPENSATION (R)
RUDDER RATIO CHANGER LINKAGE RUD
MODULE
YAW DAMPER
SERVOS

AIL ELEV AIL

POSITION
TRANSMITTER
EICAS DISPLAY
RATIO CHANGER STATUS PAGE
ACTUATOR

A RATIO
TRIM ACTUATOR CHANGER
MECHANISM
RUDDER PEDALS
FWD
A
AFT QUADRANT
ASSEMBLY RUDDER TRIM SWITCH
FEEL, CENTERING, AND AND TRIM POSITION
TRIM MECHANISM INDICATOR

FLIGHT CONTROL DIRECTIONAL AUTOPILOT SERVOS (3)

COMPUTERS L,C,R

FLT CONT/HYD - RUDDER CONTROL SYSTEM

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YSM - YAW DAMPER INTRODUCTION

Dutch Roll

Dutch Roll is a common oscillatory condition caused by low-drag, high-speed

aerodynamic design and turbulence that is created by air mass instability.
The yaw damper system gives damping with the rudder to decrease dutch roll
by the measurement of lateral acceleration and yaw rate.

Turn Coordination

The yaw damper system gives additional rudder commands to prevent yaw
when the airplane rolls into a turn. The roll induced yaw is opposite to the
intended turn direction. Rudder deflection is necessary to have a coordinated
turn.
 SIDESLIPS TO RIGHT, CYCLE REPEATS

FLIGHT
PATH BANK STARTS TURN RIGHT

WEATHERCOCKS PAST STABLE HDG

ROLLS RIGHT

HEADING MOVES LEFT, ROLLS RIGHT

(DUE TO YAW)
RELATIVE

LEFT WING INCREASES LIFT & DRAG

WIND

SIDESLIPS LEFT
CL
TURN BANK STARTS TURN LEFT
WITHOUT
RUDDER
APPLICATION
YAW WEATHERVANE
MOMENT RIGHT WING INCREASES LIFT, DRAG
ANGLE ROLLS LEFT, YAWS RIGHT
WEATHER COCKS PAST STABLE HDG

A/C MOVES RIGHT OF TRACK

TAIL MOVES RIGHT OF AXIS
SIDESLIPPING OCCURS
GUST
SIDESLIP SIDE DUTCH ROLL
FORCE

YSM - YAW DAMPER INTRODUCTION

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YSM - YAW DAMPER - GENERAL DESCRIPTION

General

The yaw damper system has two self-monitored, limited rudder authority yaw
damper functions that operate independently. They are the same. Each yaw
damper function gives a maximum of plus or minus three degrees of rudder
authority. To move the rudder, the output from the yaw damper function of the
two YSMs mechanically add together and then add in series with pilot or
autopilot commands. The two yaw damper outputs add for a maximum of plus
or minus six degrees of rudder authority.

Major Components

The yaw damper system has these major components:

- Yaw damper control panel - flight crew control of system engagement

- Left and right YSMs - control law calculation and system fault monitor
- System input sensors
- Modal suppression accelerometers
- Left and right yaw damper servo actuator - rudder control.

Interfacing Systems

The yaw damper system has interfaces with these systems/components:

- Left and right air data computers (ADCs) - give primary and secondary air data
- Left, center, and right inertial reference units (IRUs) - give inertial data
- Flap/Stabilizer position module (FSPM) - gives stabilizer and flap position data
- Left and right flight management computers (FMCs) - give gross weight data
- Left, center, and right flight control computers (FCCs) - give engine out
moments and multi-channel engagement data.
 TEST
FCC

YAW DMPR
L

YAW DAMPER EICAS

TEST SWITCH

FMC

YAW DAMPER
L R

HYD SYSTEM
FSPM PRESS SWS
YAW DAMPER
CONTROL PANEL YAW DAMPER/STABILIZER AIR
TRIM MODULE (YSM)
YAW DAMPER
SERVOS
GND
AIR/GND
SYS

AD IRU ACCELEROMETERS

YSM - YAW DAMPER - GENERAL DESCRIPTION

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YSM - YAW DAMPER - COMPONENT LOCATION

Flight Compartment

The left and right YSM and CSEU circuit breakers are on the P11 overhead
circuit breaker panel.

The yaw damper control panel is on the P5 overhead panel.

The yaw damper test switch is on the P61 right side panel.

Main Equipment Center

The yaw damper/stabilizer trim modules are on the E1-1 and E2-1 shelves.

Vertical Stabilizer

Two yaw damper servos are in the rear spar of the vertical stabilizer. Access is
through the trailing edge service access panel of the vertical stabilizer.

Aft Cargo Compartment

Two modal suppression (M/S) accelerometers are in the ceiling of the aft cargo
compartment. Access to these accelerometers is through the aft cargo door.
 P5 OVERHEAD PANEL P11 OVERHEAD CIRCUIT BREAKER PANEL
- YAW DAMPER CONTROL PANEL - LEFT YDM CB
- RIGHT YDM CB
- LEFT AND RIGHT FLT CONT
ELEC PWR SUPPLY C/B'S
MECHANICAL
LINKAGE
(REF)
RUDDER
YAW DAMPER POWER
SERVOS CONTROL
ACTUATORS
(REF)

FORWARD OF AFT BULKHEAD

- 2 MODAL SUPPRESSION
ACCELEROMETERS
P61 RIGHT SIDE PANEL
- YAW DAMPER TEST SWITCH

E2-1 SHELF
- RIGHT YAW DAMPER MODULE
- FLT CONT ELEC PWR SUPPLY MODULES (2)
E1-1 SHELF
- LEFT YAW DAMPER MODULE
- FLT CONT ELEC PWR SUPPLY
MODULES (2)

YSM - YAW DAMPER - COMPONENT LOCATION

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YSM - GENERAL DESCRIPTION Hydraulic and Air/Ground Inputs

General All the CSEU modules receive hydraulic pressure switch and air ground relay
signals for control, test, and fault indication functions.
The control system electronic unit (CSEU) modules do control functions in the
elevator, spoiler, stabilizer, and rudder systems. There are two yaw damper/ BITE Functions
stabilizer trim modules (YSMs). The YSMs use different inputs to give outputs to
actuators and modules. The YSM built-in test equipment (BITE) software give fault messages and
details to identify yaw damper and stabilizer trim system component faults.
Stabilizer Trim Function

The YSM gives manual and automatic stabilizer trim commands, mach trim, and
unscheduled stabilizer detection and annunciation. The YSMs receive stabilizer
trim control inputs from these components:

- FCCs
- Alternate electric stab trim switches on the control stand
- Manual electric trim switches on the control wheels
- Stabilizer position from a flap/slat position module (FSPM).

The YSM gives control signals to a stabilizer trim control module (STCM)
through the cutout and limit switches.

The air data/inertial reference units (ADIRU’s) give mach and computed
airspeed for the mach trim function.

Caution, advisory, and maintenance messages go to the EICAS computers and

pilots overhead panel for display.

Yaw Damper Function

The YSM receives inputs from these components to control the yaw damper
actuator:

- Air Data Inertial Reference Units (ADIRU)

- Flight management computers (FMCs)
- Accelerometers
- Position signals from the yaw damper actuator LVDT
- FCC.
 MANUAL TRIM

TEST
ALTERNATE TRIM
EICAS

YAW DMPR
L
STCM
R CUTOUT LIMIT
SWITCHES SWITCHES
YAW DAMPER
TEST SWITCH
FCC

P5 PANEL

FMC
YAW DAMPER/STABILIZER
FSPM TRIM MODULE (YSM)

YAW DAMPER HYD SYSTEM

L R
PRESS SWS
ADIRU (3)
AIR YAW DAMPER
YAW DAMPER SERVOS
CONTROL PNL
GND
AIR/GND SYS
ACCELEROMETERS

YSM - GENERAL DESCRIPTION

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YSM - GENERAL DESCRIPTION - FUNCTIONS

General

The YSM is part of the control system for these flight control surfaces:

- Rudder
- Horizontal stabilizer
- Spoilers
- Elevators
- Ailerons

There are seven functions that are done by the YSM.

 BETA DOT YAW DAMPING CALCULATION OF VC AND MC LIMITS
STRUCTURAL MODA L SUPPRESSION
STRUCTURAL NOTCH FILTER CONTROL WHEEL INHIBIT TO SPOILER CONTROL MODULE
YAW RATE DAMPING AT HIGH ANGLE OF ATTACK REDUNDANCY MANAGEMENT
SYSTEM D ISENGAGEMENT FAILURE INDICATION
FAILURE INDICATIONS
MAINTENANCE INDICATIONS OUTBOARD AILERON LOCKOUT FUNCTION
YAW DAMPER FUNCTION

RUDDER COMMAND GAIN CONTROL LAW CALCULATION OF ELEVATOR FEEL TAKEOFF PRESSURE
REDUNDANCY MANAGEMENT LIMIT
FAILURE INDICATION FAILURE INDICATION
MAINTENANCE INDICATIONS ELEVATOR FEEL LIMIT FUNCTION
RUDDER RATIOCHANGER FUNCTION

NON-VOLATILE STORAGE AND RETRIEVALOF FAILURE AND

AUTOTRIM INTERFACE MAINTENANCE RELATED DATA
MANUAL ELECTRIC TRIM INTERFACE SELF-TEST
MACH TRIM CONTROL LAW MAINTENANCE TEST
STABILIZER TRIM MODE SELECTION
MANUAL TRIMRATE MONITORING BITE FUNCTION
UNSCHEDULED TRIM DETECTION
REDUNDANCY MANAGEMENT
FAILURE INDICATION
MAINTENANCE INDICATIONS
STABILIZER T RIM FUNCTION

AUTO SPEEDBRAKE ENA BLE CALCULATION

SPEEDBRAKE THETA INHIBIT CALCULATION
FAILURE INDICATION
SPEEDBRAKE FUNCTION

YAW DAMPER/STABILIZER TRIM MODULE

YSM - GENERAL DESCRIPTION - FUNCTIONS

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YSM - FAULT RECORDING Flight Leg

A new flight leg starts on the transition to takeoff. At takeoff, the faults history
General
flight leg(s) will increment. The current flight leg is shown as flight leg zero.
The YSM contains non-volatile memory (NVM) to record fault data. There is
Fault Record
memory to store 256 faults and 64 flight legs.
A fault record is stored for each flight leg in which a fault is recorded. The fault
Fault Storage
record(s) for each existing fault and ground test fault may contain this data:
Because the EICAS receives discretes from the YSMs that are connected
- Fault message
together, the YSMs are the only place to find fault details and identify which
- Message number
YSM detected the
- Flight deck effect (FDE)
fault. Faults can be stored from any of these:
- Most likely line replaceable units (LRUs) at fault (shown as the reference
designator for the LRU)
- Power-up test
- More possible faulty LRUs or details that cause the fault (if any)
- Ground test
- Additional details or wiring diagram (if necessary)
- Flight phase faults
- Fault latched or not (not shown in fault history).
- Internal YSM faults.
The fault record(s) for fault history may contain the above details and may
Power-up test is done at initial power-up of the YSM. The test lasts no more
include this additional data:
than five seconds. The yaw damper INOP amber light is turned on and an
EICAS message L/R YAW DAMPER shows during the power-up test. If a fault is
- Fault type (hard, intermittent, or repeated)
detected, the INOP light and EICAS message stay.
- Flight phase of first occurrence.
Ground test can be started when the airplane is on the ground and the ground
speed is less than 60 knots. The flight phase faults start at takeoff after 60 knots
ground speed and ends at landing when the ground speed is less than 60 knots.
These are the seven flight phases:

- Takeoff
- Climb
- Cruise
- Holding
- Descent
- Approach
- Landing.

All internal YSM faults are stored regardless of when they occur. Internal faults
are identified by the fault message YSM FAULT.
 FAULTS HISTORY
FLIGHT LEG = X FLIGHT LEG = 0

DESCENT

CRUISE HOLDING
EXISTING FAULTS
CLIMB FAULT HISTORY
GROUND TEST
POWER-UP TEST APPROACH
GROUND TEST TAKEOFF FAULT MSG NO: ON GROUND
MESSAGE XX-YYY LANDING

GROUND SPEED
> 60 KNOTS FDE
(TEXT)
EXISTING
SER
-CAUTION-
FAULTS?
MFR
REMOVE THE FRONT PANEL
SLOWLY TO PREVENT DAMAGE
TO INTERNAL CABLE
LRU:
TEXT/*
FAULTS
HISTORY? (MORE POSSIBLE
MENU ON/OFF
FAULT LRUS OR
YES NO
DETAILS)
GROUND PERFORM FAULT HISTORY ONLY
TEST? GROUND (ADDITIONAL
TEST DETAILS, WIRING
YAW DAMPER/STABILIZER TRIM
MODULE (YSM)
P/N 285T1122-
DIAGRAMS) TYPE:
SERIAL NUMBER
MOD D C B A
LEVEL
H G F E
OTHER EXISTING FAULTS
FUNCTIONS? GROUND TEST ONLY
PHASE:
LATCHED:
(YES OR NO)

YSM - FAULT RECORDING

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YSM - BITE

General

The YSM built-in test equipment (BITE) software provides fault messages and
details to identify yaw damper and stabilizer trim system component faults.
Entry and exit for BITE is through the BITE control module (BCM) on the front of
the YSM. Instructions for the use of the buttons on the BCM are on a placard
attached to the front of the YSM.

LRU Reset is available by pressing both the YES and NO buttons at the same
time. This performs a maintenance reset which includes a software reset to
reset the latched messages.
 INSTRUCTIONS:
Push ON/OFF to start or stop BITE display
Push YES or NO in reply to questions(?)
CAUTION Push to move down in list
OBSERVE PRECAUTIONS
FOR HANDLING Push to move up in list
ELECTROSTATIC
SENSITIVE
INSTRUCTIONS: Push MENU to return to previous menu
DEVICES

BITE MAIN MENU:

BITE MAIN MENU: EXISTING FAULTS - Sh ows existing faults
FAULT HISTORY - Show s previous faults by flight leg
GROUND TESTS - Shows lists of ground tests
LRU RESET: OTHER FUNCTIONS - Shows other functions

LRU RESET:
Push YES and NO simultaneously for 1 second

EXISTING
YAW DAMPER/STABILIZER FAULTS?
TRIM MODULE
P/N 285T1122-
S/N
MOD LEVEL A B C D E

YAW DAMPER/STABILIZER TRIM MODULE ON

MENU
OFF

YES NO

YSM - BITE
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YSM - FAULT ANNUNCIATION - Accelerometer fault

- Yaw damper acceleration fault (excitation)
General - FMC fault
- Reference 26v ac fault.
Flight deck effects (FDEs) for the YSM show on these:
The message, STAB TRIM (M) shows for any of these:
- Upper and lower EICAS displays
- P5 overhead panel - Channel pin fault
- Yaw damper control panel. - 15 volt power fault
- Stab trim control module (STCM) fault
Upper EICAS Display - Left and center flight control computer (FCC) (left YSM)
- Right and center FCC (Right YSM)
The message, UNSCHD STAB TRIM (B) shows for any of these: - Stabilizer position
- Flap 1, 2, and 3 fault
- Stabilizer movement is detected without a valid stabilizer trim command from - Manual trim switch fault
the YSM - left hydraulic pressure fault (Left YSM)
- Stabilizer movement opposite to a valid stabilizer trim command from the YSM - Center hydraulic pressure fault (Right YSM)
- Alternate electric trim is used when an autopilot is engaged. - Stab brake applied
- Stab 28v dc fault
The message, STAB TRIM (C) shows if both YSMs command manual trim and - YSM fault (internal fault).
only one brake releases.
The message, YSM (S) shows for a YSM internal fault.
The message, L/R YAW DAMPER (C) shows for any of these:
P5 Overhead Panel
- Channel pin fault
- Airplane pin fault The lights on the P5 panel come on for the same reasons as the EICAS
- Invalid YSM part number messages.
- Yaw damper actuator fault
- Yaw damper intermittent Yaw Damper Control Panel
- YSM fault (internal fault)
- L/R yaw damper disengaged. The yaw damper INOP lights on the yaw damper control panel come on for the
same reasons as the L/R YAW DAMPER (C) message.
Lower EICAS Display

The message, YAW DAMPER (M) shows for any of these:

- Yaw damper LVDT fault

- Left hydraulic pressure fault (Right YSM)
- Center hydraulic pressure fault (Left YSM)
- Left, center, right IRU faults
- YSM fault
 YAW DAMPER
L R

ON w
ON w

INOP INOP TAT-12C

a a

CRZ
1.50 1.50
YAW DAMPER UNSCHD STAB TRIM
1.8
1.00
1.8
1.00
CONTROL PANEL STAB TRIM
L YAW DAMPER 1.0 1.0
1.4 1.4
R YAW DAMPER

00.0 00.0
WARNING
r

CAUTION N
1
a
21 21
MASTER CAUTION/
WARNING LIGHT (2)
GET

YSM

YAW DAMPER
STAB TRIM

STAB
TRIM a
STATUS DISPLAY
UNSCHED
STAB TRIM
a

P5 PANEL ECS/MSG DISPLAY

YSM - FAULT ANNUNCIATION

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YSM - YAW DAMPER - PRE FLIGHT TEST FROM THE FLIGHT The IRUs transmit these normal IRU test values over the ARINC 429 bus:
DECK - Pitch angle
- Roll angle
General
- Yaw rate
- Roll rate
The preflight test uses only the built-in test functions of the yaw damper
- Lateral acceleration
function. To start the preflight test, push the YAW DMPR test switch on the P61
- Ground speed.
panel. The test makes sure that the yaw damper functions of the yaw damper/
stabilizer trim module (YSM) operate correctly. It also makes sure the yaw
Because the YSM is in the preflight test mode, it uses these IRU test values and
damper servos and the switches on the yaw damper control panel operate
monitors them for correct performance. If an IRU input is not in limits, an L/C/R
correctly.
IRU fault message shows.
Hydraulics
The rudder position moves 3 degrees right, goes back to 0, then moves 3
degrees left, and goes back to 0. The test is complete in fifteen seconds. The
The center hydraulic system is necessary to do a test of the left yaw damper.
yaw damper INOP annunciator goes off approximately fifteen seconds after test
The left hydraulic system is necessary to do a test of the right yaw damper.
start to show satisfactory
completion of the test.
Interfacing Systems
The mode and status annunciators on the inertial reference mode panel come
These systems are necessary during the yaw damper test:
on during the IRU test mode.
- Rudder and rudder trim control system
IRS test values can also show on flight compartment indications. See the air
- IRU aligned and in NAV mode
data inertial reference unit section for more information about the IRS test (ATA
- Air/ground relays
34-21).
- Master dim and test system
- EICAS.
If at least one IRU data is satisfactory, the INOP indications stop, and the
system goes back to the ground operation mode. If all three IRU data tests have
Yaw Damper Systems
failures, the INOP indication stays on, and the auto disengage relay de-
energizes.
Push the YES and NO button on the left or right YSM BITE control module
(BCM) to do a reset of the YSMs. After 30 seconds, push the left or right yaw
BCM Preflight Test Indications
damper engage switch on the yaw damper control panel and make sure the ON
light comes on, and the INOP light goes off.
When a preflight test starts from the flight compartment, the BCM shows YD
SYS IN TEST. This lets the person in the main equipment center know there
Operation
was a preflight test started from the flight compartment.
Momentarily push the yaw damper test switch up or down to start the preflight
test in the left or right yaw damper module. The INOP annunciator for the
WARNING: KEEP PERSONNEL AND EQUIPMENT AWAY FROM ALL
channel in test is on during test of each channel. The yaw damper module starts CONTROL SURFACES WHEN HYDRAULIC POWER IS
a test in all three IRUs. SUPPLIED. INJURY TO A PERSON OR DAMAGE TO
EQUIPMENT CAN OCCUR WHEN HYDRAULIC POWER IS
SUPPLIED.
 PRE FLT ENGAGE
YD TEST? YD

TURN ON
MENU
ON
OFF
HYD

YES NO
WARNING!
WARNING! 3O 3O
L R
C
RUDDER CLEAR
CLEARED? RUDDER

YD SYS PASS/FDE FMC

IN TEST DUE TO: FAULT

CHANNEL TESTED
INOP LIGHT ON FOR TEST
TIME OF PREFLIGHT TEST PASS
YAW DAMPER
L R RUDDER RUD
INDICATOR
ON w
ON w
INOP a INOP a AIL ELEV AIL

YAW DAMPER CONTROL PANEL EICAS STATUS PAGE

YSM - YAW DAMPER - PRE FLIGHT TEST FROM THE FLIGHT DECK
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YSM - STABILIZER TRIM - INTRODUCTION

Purpose

The horizontal stabilizer is a moveable airfoil that adjusts the airplane trim in the
pitch axis. As the aerodynamic characteristics of the airplane change, the
angle of attack of the stabilizer changes to hold the airplane longitudinal
stability.

Movement of the stabilizer leading edge up gives the airplane nose down trim.
Movement of the stabilizer leading edge down gives the airplane nose up trim.
Stabilizer trim is done by these trim modes:

- Manual electric or alternate electric stabilizer trim switches

- the pilot gives stabilizer trim commands
- Auto trim mode - an engaged autopilot (FCC) gives stabilizer trim commands
- Mach trim mode - retraction of the flaps cause stabilizer trim as Mach
increases.

Automatic stabilizer trim and Mach trim modes have a relation to the manual
stabilizer modes and use common system components.
 STABILIZER
TRIM CONTROL
MODULE (2)

APL
CONTROL NOSE DN
SWITCH ARM
SWITCH

APL
NOSE UP

ALTERNATE
ELECTRIC
STAB TRIM LIMIT SWITCH
AND POSITION
SWITCHES TRANSMITTER
TO STCM MODULES
STABILIZER TRIM BALLSCREW
ACTUATOR ASSEMBLY
STABILIZER
POSITION
MANUAL ELECTRIC INDICATOR (2)
TRIM SWITCHES
ADC FAIL
COLLINS
EXTERNAL TEST

MACH DATA
SENSOR ONLY
FAULT
FUNCTIONAL
TEST

TRIM
COMMANDS
ADC (2) TO STCM
THROUGH
FLIGHT CONTROL AUTOTRIM DATA LIMIT AND
COMPUTER (3) CUT-OUT
MANUAL ELECTRIC TRIM CMD SWITCHES

CAPT CONTROL F/O CONTROL YAW DAMPER/STABILIZER

WHEEL (OUTBD) WHEEL (OUTBD) TRIM MODULE (2)

YSM - STABILIZER TRIM - INTRODUCTION

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YSM - STABILIZER TRIM - COMPONENT LOCATION - FLIGHT Stabilizer Trim Control Switches
DECK The stabilizer trim control switches send manual electric arm and control
commands to the YSMs. Each set of switches on the captain and first officer
Control Column Cutout Switches control wheel has an arm and a control switch. It is a three position rocker
switch spring-loaded to the center (off) position. One or the other set of switches
The column cutout switches give manual override of an electric trim command moved down gives an airplane nose up trim command, and one or the other set
when the control columns move in a direction opposite to the electric trim moved up gives an airplane nose down trim command.
command. This corrects a runaway trim.
Alternate Electric Stab Trim Switches
P5 Overhead Panel
Two switches (arm and control) on the P10 quadrant stand give electrical input
The STAB TRIM light shows if the two YSMs give a manual trim command, and to the STCMs.
only one brake releases. The UNSCHD STAB TRIM light shows for these
conditions: Training Information Point

- Stabilizer movement without a valid stabilizer trim command from the YSM Access to the control column cutout switches is through the forward equipment
- Stabilizer movement opposite to a valid stabilizer trim command from the YSM center. The other items are in the flight compartment.
- Alternate electric trim operated when an autopilot is engaged.

P1-3 Captain Instrument Panel

The autopilot caution light shows a loss of stabilizer trim control from the active
autopilot.

Stabilizer Position Indicators

Vertical scale instruments show the stabilizer position. Zero units is equal to the
four degrees stabilizer leading edge up mechanical stop. The stabilizer leading
edge down mechanical stop is 15.5 units (-11.5 degrees). The green band (2.0
to 7.0 units) shows the takeoff range. The OFF flag shows a black and amber
striped area.

Stabilizer Trim Cutout Switches

The guarded switches control 28v dc standby bus power to the STCM hydraulic
cutoff valves. The switch on the left (L) removes left system hydraulics to the left
STCM. The switch on the right (C) removes center system hydraulics to the
right STCM. The guarded switches are usually in the NORM position (open
cutoff valve)

.
 ENTRY EMER CARGO ACCESS
DOORS a DOORS a DOORS a DOORS a

CAPT FO
PITOT a PITOT L AOA R AOA
a a a
L AUX R AUX
L TAT R TAT
MANUAL PITOT a PITOT a a w
ELECTRIC STAB AUTO
CONTROL TRIM a
SPOILERS
a SPDBRK a
SWITCHES UNSCHED RUDDER
ANTISKID
AIL
STAB TRIM a RATIO a a LOCK a

BODY L ENG R ENG

VANE a w PROBE a PROBE a

w w w w

CONTROL COL
CONTROL COL CUTOUT SW ANNUNCIATOR PANEL (P5)
CUTOUT SW (UNDER FLOOR
(UNDER FLOOR FIRST OFFICER
CAPTAIN CONTROL COL)
CONTROL COL)
APL
NOSE 0
DN
ALTERNATE S 2
FIRE CONFIG
r r STAB TRIM T 4
SWITCHES A
AUTOPILOT A/P B 6
CAUTION PULL UP
r DISC r 8
LIGHT T 10
TAPE
CABIN
OVSPD
R
ALT r r I 12
M 14
ALT AUTO NORM APL
ALERT a PILOT a NOSE
GND QUADRANT STAND UP
a
A/T PROX G/S
(P10)
FMC INHB a
DISC a a
CUT
P1-3 CAPT INST PANEL OUT
STABILIZER
POSITION
L C
INDICATOR (2
(2)
STAB TRIM
STAB TRIM CUTOUT SW

YSM - STABILIZER TRIM - COMPONENT LOCATION - FLIGHT DECK

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 B767-3S2F ATA 22-00 TRAINING MANUAL
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YSM - STABILIZER TRIM - COMPONENT LOCATION - MEC

AND JACKSCREW AREA
Main Equipment Center

Yaw damper/stabilizer trim modules (YSMs) are CSEU modules on the CSEU
shelf.

Stabilizer Jackscrew Access Area

The stabilizer trim actuator controls trim operation of the horizontal stabilizer.
Each trim limit switch and position transmitter module (left, center, and right)
contains these components:

- One synchro
- One rotary variable differential transformer
(RVDT).

All three modules are interchangeable.

 STABILIZER TRIM
CONTROL MODULES
(LEFT AND RIGHT)

STABILIZER TRIM ACTUATOR

- HYDRAULIC BRAKES
(LEFT AND RIGHT)
- HYDRAULIC TRIM MOTORS
(LEFT AND RIGHT)
E2-1 SHELF E1-1 SHELF
- RIGHT YSM - LEFT YSM

TRIM LIMIT SWITCH

AND POSITION
TRANSMITTER
MODULES

MAIN EQUIPMENT CENTER STABILIZER JACKSCREW ACCESS AREA

YSM - STABILIZER TRIM - COMPONENT LOCATION - MEC AND JACKSCREW AREA

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 B767-3S2F ATA 22-00 TRAINING MANUAL
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YSM - STABILIZER TRIM - GENERAL DESCRIPTION Autopilot Trim System

The three FCCs in the main equipment center supply autopilot input signals to
General
the two YSMs to do a trim of the stabilizer related to elevator out of neutral
position. The function of the YSM, STCMs, stabilizer trim ballscrew actuator
The yaw damper/stabilizer trim modules (YSMs) give these trim commands to
assembly, and limit switch and position transmitter modules is the same as the
the stabilizer trim control modules (STCMs):
manual electric system.
Mach/Speed Trim System
- Manual trim
- Autopilot trim
One or the other YSM uses signals from the air data computers (ADC) and
- Mach/speed trim.
inertial reference units (IRUs) to do a trim of the stabilizer for mach/airspeed.
The FSPMs send signal to the YSM for Mach (flaps retracted) selection. High
Manual Electric Trim System
alpha speed trim is enabled when flaps are less than 5. Air/ground logic from
the air/ground relays inhibits Mach trim on the ground. The function of the
The stabilizer trim control switches on the outboard horn of each control wheel
STCMs, the stabilizer trim ballscrew actuator assembly, and the limit switch and
give trim up or down arm and control signals to the left and right YSM. The
position transmitter modules is the same as the manual electric system.
YSMs supply trim up or down arm and control signals to solenoids on the
STCMs in the stabilizer compartment. The STCM arm and control hydraulic
Alternate Electric Stab Trim System
valves operate in series to let hydraulic flow go to the two hydraulically released
brakes and the two hydraulic motors on the stabilizer trim ballscrew actuator
The two alternate electric stab trim switches give up or down arm and control
assembly forward of the stabilizer. Left and center hydraulic systems supply
trim signals to the dual coil solenoid valves in the STCM. These signals
pressure to the STCMs through STCM cutoff valves. Two hydraulic cutout
energize one coil in the dual coil solenoid valves (the second coil is controlled
switches on the top left side of the control stand supply electrical power to the
by the trim signals from the YSMs). The valves operate as before and supply
STCM cutoff valve motors. This isolates hydraulic power from the STCMs.
hydraulic pressure to release the hydraulic brakes and move the hydraulic
Hydraulic inputs from the elevator feel computers control the rate of stabilizer
motors to do a trim of the stabilizer.
trim. The left and right limit switch and position transmitter modules near the
stabilizer supply position feedback through flap/stab position module (FSPM) to
Stabilizer Trim Fault Indication System
the flight control computers (FCCs). The position and limit switch modules also
control the flight compartment stabilizer position indicators and limit stabilizer
The YSMs and FCCs control the logic for the flight compartment amber
travel by cam-operated microswitches that open the electrical path from the
annunciation on the P5 overhead panel and the indication on EICAS. The YSM
YSMs to the STCMs. Column-operated cutout switches below the floor on each
also has a BITE control module (BCM) to make a record of the status of
outboard side of the control column torque tube also open the electrical path
system LRUs.
from the YSMs to the STCMs when the control columns move in a direction
opposite to stabilizer trim. Trim limit select relays control selection of stabilizer
Stabilizer Trim Ground Test
electrical trim limits related to trailing edge flap retract signal from the FSPM.
The maintenance control and display panel (MCDP) does a test of the ability of
each FCC (L, C, R) to connect with the YSMs and move the horizontal stabilizer
up or down.
 YSM - STABILIZER TRIM - GENERAL DESCRIPTION
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 B767-3S2F ATA 22-00 TRAINING MANUAL
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YSM - STABILIZER TRIM - FUNCTIONAL MODE PRIORITY Manual electric trim inhibits during multi-channel operation. The YSM does not
let the manual electric trim inputs disengage the autopilot.
General

There are four modes for control of stabilizer trim. The mode priority has a
relation to the number of autopilot channels engaged. The four modes are:

- Alternate electric trim mode

- Manual electric trim mode
- Mach trim mode
- Automatic stabilizer trim mode.

Autopilot Disengaged

Mach trim enables when the autopilot is not engaged in CMD and alternate
electric or manual electric trim is not in use. This occurs automatically and mode
selection is not necessary. Operation of the manual electric trim or the alternate
electric stab trim mode causes an interrupt of Mach/speed trim commands.

Single Channel Autopilot Engaged

When the autopilot engages in single channel, this is the order of control
priority:

- Alternate electric trim mode

- Manual electric trim mode
- Automatic stabilizer trim mode.

Operation of manual electric trim causes the autotrim valid signal to be set low
from the YSM. This causes the FCC to disengage during single channel
operation. Operation of alternate electric trim does not disengage the autopilot.

Multi-Channel Autopilot Engaged

When the autopilot is multi-channel engaged (two or more FCCs engaged), this
is the order of priority:

- Alternate electric trim mode

- Automatic stabilizer trim mode.
 AUTOPILOT DISENGAGED SINGLE CHANNEL AUTOPILOT MULTICHANNEL AUTOPILOT
ENGAGED ENGAGED

1 ALTERNATE ELECTRIC 1 ALTERNATE ELECTRIC 1 ALTERNATE ELECTRIC

OR OR STAB TRIM
MANUAL ELECTRIC MANUAL ELECTRIC
STAB TRIM STAB TRIM 2 AUTOSTAB
TRIM
2 MACH 2 AUTOSTAB
TRIM

(OPERATION OF MANUAL ELECTRIC (MANUAL ELECT RIC TRIM

TRIM DISENGAGES AUTOPILOT) INHIBITED IN MULTICHANNEL
ENGAGE)

YSM - STABILIZER TRIM - FUNCTIONAL MODE PRIORITY

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 B767-3S2F ATA 22-00 TRAINING MANUAL
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STABILIZER TRIM SYSTEM - MANUAL ELECTRIC STAB TRIM

- FUNCTIONAL DESCRIPTION
Manual Electric Control Mode

Operation of the captain or first officer control wheel switches gives full trim rate
to the horizontal stabilizer (approximately 0.5/sec on the ground). The STCMs
supply hydraulic power to release brakes and hydraulic motors to move the
stabilizer through the stabilizer trim ball screw actuator assembly. The arm
portion of the signal goes through column cutout and limit switches which open
the trim signal to STCM solenoids. This stops stabilizer trim.

The stabilizer trim column cutout switches give manual override of any electric
trim command if there is a runaway stabilizer. It breaks the path for electric
trim commands when you move the control columns in a direction opposite to
the commanded trim direction. The normally closed switches are in parallel to
give column asymmetry protection. If you move the columns forward
more than the limit, it stops a stabilizer airplane nose up command. If you move
the columns aft more than the limit, it stops a stabilizer airplane nose down
command.

Two switches in the limit switch and position transmitter modules limit the travel
of the stabilizer in a leading edge up direction (airplane nose down
trim). The leading edge up electrical limit is 1.5 units for flaps retracted and 0.25
units for flaps not retracted.
 CLOSE S/O
STAB TRIM VALVE
28V DC CUTOUT CONTROL NOSE UP CONT
STBY BUS NORM SOLENOIDS
L STAB TRIM ARM
SHUTOFF NOSE UP ARM
CUTOUT SOLENOIDS
28V DC A FWD
STBY BUS OPEN S/O
NORM L COL FWD VALVE
R STAB TRIM B NOSE
SHUTOFF UP NOSE DN CONT
FLAPS UP SOLENOIDS
P11 CB PANEL P10
FWD NOSE DN ARM
SOLENOIDS
NOSE R COL FWD
L TRIM
DN ARM LIMIT
SELECT SECONDARY
BRAKE
AFT RELEASE
NOSE UP R COL AFT L
ARM STCM DN UP
28V DC
STBY HYD HYD
BUS AFT BRAKE MOTOR
CONTROL LIMIT SWITCHES
L STAB L COL AFT
NOSE DN POSITION STABILIZER
TRIM CONT
CAPT CONT L YSM XDUCER BALLSCREW
WHEEL TRIM (E1-1) L POSN XMTR ACTUATOR
SWITCH AFT MODULE ASSEMBLY
28V AC
L BUS L COL AFT L STAB
COLUMN CUTOUT POS IND HYD HYD
L CONT
STAB TRIM SWITCHES P10 BRAKE MOTOR
P11 CB PANEL
SAME ARM SAME DN UP
AS AS (RIGHT CHANNEL SAME AS LEFT)
CAPT LEFT A SAME
CONTROL AS
F/O CONT R YSM B LEFT
WHEEL TRIM (E2-1)
SWITCH R STCM

STABILIZER TRIM SYSTEM - MANUAL ELECTRIC STAB TRIM - FUNCTIONAL DESCRIPTION

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 B767-3S2F ATA 22-00 TRAINING MANUAL
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STABILIZER TRIM SYSTEM - ALTERNATE ELECTRIC STAB

TRIM - FUNCTIONAL DESCRIPTION
Alternate Electric Stab Trim

Operation of the alternate electric stab trim switches gives full rate trim.
Movement of the arm (right) and the control (left) switches forward or aft gives a
nose up or nose down signal. When you release the switches, a spring moves
the switches back to the neutral position.

The nose up or down arm and control signals go directly to the two STCMs.

A normally open micro-switch closes with arm switch movement to connect a

ground to the stabilizer trim aileron lockout modules (SAM). This inhibits the
Mach trim function. The SAM monitors the micro-switch position. If the switch is
in the closed position for more than 30 seconds, and one of the systems has
hydraulic pressure, the manual lever switch fault ball sets on the SAMs.
 ARM NOSE DN CONTROL
SWITCH FLAP ALT STAB SOLENOIDS
SW FAULT NOSE UP CONTROL
SOLENOIDS
CONTROL L YSM NOSE UP ARM
SWITCH APL
NOSE DN (E1-1) SOLENOIDS
NOSE DN ARM
APL
NOSE UP
SOLENOIDS
ALTN
STAB TRIM ENG ENG
VALVE a
SPAR
VALVE a
REV
ISLN a
VALVE a
SPAR
VALVE a
SECONDARY
L FUEL CONTROL
RUN
R
BRAKE
NORM
CUT OFF
ALT STAB RELEASE
CUT SW FAULT
L STCM
OUT

L
STAB TRIM
C
DN UP
R YSM HYD HYD
(E2-1) BRAKE MOTOR
STABILIZER
BALLSCREW
ACTUATOR
UP ASSEMBLY
IND
DN HYD HYD
BRAKE MOTOR
DN UP
28V DC UP
BAT BUS ARM SECONDARY
DN
BRAKE
ALT STAB RELEASE
TRIM UP
CONTROL NOSE UP CONTROL
P11 CB PANEL DN SOLENOIDS
NOSE UP ARM
SOLENOIDS
ALTN ELEC STAB NOSE DN ARM
TRIM SWITCHS SOLENOIDS
NOSE DN CONTROL
SOLENOIDS
R STCM

STABILIZER TRIM SYSTEM - ALTERNATE ELECTRIC STAB TRIM - FUNCTIONAL DESCRIPTION

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 B767-3S2F ATA 22-00 TRAINING MANUAL
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YSM - STABILIZER TRIM - FUNCTIONAL DESCRIPTION -

AUTOMATIC STABILIZER TRIM
General Description

Automatic trim of the stabilizer is done by trim up and trim down commands
from an engaged FCC. The first FCC engaged starts communications with its
YSM. The left and right FCCs interface with the left and right YSMs respectively.
The center FCC interfaces with both YSMs. Each FCC has logic that
determines which YSM processesthe FCC autotrim commands. The first
engaged FCC provides the autotrim signals and controls YSM selection. When
multichannel engaged, if the first FCC in command disengages or is unable to
provide reliable autotrim commands, an automatic change to another FCC
occurs. Autotrim failure annunciation is controlled by fault logic in the FCC in
control.

Functional Description

Each YSM makes an autotrim valid discrete if it is capable of the trim function.

The FCC sends a discrete autotrim arm signal to the YSM(s). This arms the
YSM(s) for data reception. The FCC then sends an engage discrete and
autotrim commands on the FCC-YSM 429 bus interface based on elevator out
of neutral position. The YSM processes the FCC autotrim commands through
arm and control microprocessors to provide up and down signals to the stab
trim control module (STCM).
 MANUAL COMMANDS
SPEED TRIM COMMANDS
TO ENGAGE
INTERLOCKS AUTOTRIM VALID SRM FAULT
AND AUTOTRIM D
A VALID MANUAL COMMAND
LOGIC TO MCP
T
1 A
AUTOTRIM ARM
M
LEFT FCC (E1-1) A
TO MCP N
A
G COMMANDS
E 3.5 SEC TO STCM
SAME AS M AUTOTRIM
LEFT FCC E COMMANDS
N FRAT
T TRIM
CENTER FCC (E1-3) TO MCP TRIM
TRANSFER
LOGIC TRANSFER
AUTOTRIM SELECTION RELAY
LOGIC
SAME AS
LEFT FCC LEFT YSM (E1-2)
SAME AS LEFT SRM

RIGHT FCC (E1-4) RIGHT YSM (E2-2)

1 ASTS DATA
- AUTOTRIM ENGAGE LEFT/RIGHT
- TRIM UP/DOWN ARM
- TRIM UP/DOWN CONTROL
- FULL RATE AUTOTRIM

YSM - STABILIZER TRIM - FUNCTIONAL DESCRIPTION - AUTOMATIC STABILIZER TRIM

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YSM - STABILIZER TRIM- FAULT ANNUNCIATION - STAB - YSM VAL DISCRETE.

TRIM Stab Brake Switch Fault
General
This fault is caused if the two YSMs give a manual trim command and only one
A YSM fault message that causes a STAB TRIM level C advisory EICAS
brake releases.
message and a STAB TRIM annunciator on the P5 overhead panel can show
on the EXISTING FAULTS? and FAULT HISTORY? menus. This message is
Alternate Stab Switch Fault
set and latched. The term latched used in BITE means that more than repair of
the fault is necessary to remove the fault indication. A software reset is
This fault is caused by one of these:
necessary.
- An alternate electric stab trim micro-switch failure
The YSM fault messages that causes STAB TRIM level M maintenance EICAS
- A Mach loop error - stabilizer movement slower than 0.067 degrees/second for
message can show on the EXISTING FAULTS? and FAULT HISTORY? menus.
10 seconds of a valid command or 0.3 degrees stabilizer position difference
between stabilizer position and commanded position for 10 seconds.
These fault messages are set and latched:

- STCM fault
- MAN TRIM sw fault
- L/C hydr discrete
- STAB BRK sw fault
- STAB 28v fault.

Do one of these to do a stab trim software reset:

- Set the +/- 5v dc control system electronic unit (CSEU) power supply off then
back on
- Pull the stab trim 28v dc circuit breaker then push it back in
- Do a RESET LATCH? procedure from the EXISTING FAULTS? menu.

These fault messages also cause the STAB TRIM level M maintenance EICAS
message and do not reset with a stab trim reset:

- ALT STAB SW FAULT

- CHNL PIN FAULT
- 15V PWR FAULT
- FCC L/R/C FAULT
- L/R/C FSPM FAULT
- YSM FAULT
- AIR/GND 1/1_2/2 FAULT
- L/R ADC FAULT
- YSM BUS XFEED
 - STCM FAULT
- MAN TRIM SW FAULT
- L/C HYDR DISCRETE
- STAB BRK SW FAULT STAB
- STAB 28V FAULT S Q TRIM
a
R PANEL
EXISTING RESET LATCH? P5 OVERHEAD
FAULTS?
CSEU +/- 5V DC

STAB TRIM 28V DC

ON
MENU OFF
- ALT STAB SW FAULT
YES NO
- CHNL PIN FAULT
- 15V PWR FAULT
- FCC L/R/C FAULT
- L/R/C FSPM FAULT
- YSM FAULT
- AIR/GND 1/1_2/2 FAULT TAT

- L/R ADC FAULT STAB TRIM

- YSM BUS XFEED WARNING
r
- YSM VAL DISCRETE
CAUTION
a

L/R YSM

EICAS

YSM - STABILIZER TRIM- FAULT ANNUNCIATION - STAB TRIM

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YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB Channel Pin Fault

TRIM CONT. The YSM channel code monitor makes sure of the validity of the channel code
discrete inputs from hardware and makes an analysis of these discretes to find
STCM Fault in which position (left or right) the YSM is installed. If there is a disagreement
between the channel code discretes, a channel code fault is set. If the channel
The STCM fault is caused by one of these: code discretes come back into agreement, the channel code fault has an auto
reset.
- Stabilizer movement is slower than 0.067 degrees/second for 10 seconds of a
valid command 15v Power Fault
- 0.3 degrees difference between stabilizer position and commanded position
for 10 seconds. If the +/-15v dc power input is not in the valid range (> +10v dc for +15v dc and
< -10v dc for -15v dc) for 2 seconds, a 15v pwr fault is set. The fault has an auto
The fault has a reset 5 seconds after the error goes away or by a stab trim reset. reset when the power input comes back in valid range.
Manual Trim Switch Fault

A manual trim switch fault is set and latched if the manual electric trim inputs to
the YSM have a disagreement for 5 seconds. The fault resets when the
inputs come back into agreement or by a stab trim reset.

Left or Center Hydraulic Discrete Fault

This fault is caused if the hydraulic pressure necessary for the stabilizer function
is not available.

Stabilizer Brake Switch Fault

This fault is for a stabilizer hydraulic brake apply fault. The brake is on the
STCM. The stabilizer trim brake apply fault is set and latched if there is a valid
trim command, but there is no stabilizer hydraulic brake release. This condition
stays for 3 seconds. The fault has an auto reset if a subsequent valid command
results in a brake release or by a stab trim reset.

STAB 28v Fault

If the stabilizer trim 28v dc is off for 2 seconds, a fault is set and latched. The
fault has an auto reset when stabilizer trim 28v dc is supplied. The application of
the stabilizer trim 28v dc is part of the stab trim reset.
 - STCM FAULT
- MAN TRIM SW FAULT
- STCM
- L/C HYDRFAULT
DISCRETE
- MANBRK
- STAB TRIMSW
SWFAULT
FAULT STAB
- L/C HYDR
- STAB DISCRETE
28V FAULT
- STAB BRK SW FAULT S Q TRIM
STAB a
- STAB 28V FAULT SR Q TRIM
PANEL a
EXISTING
FAULTS?
RESET LATCH? R PANEL OVERHEAD
P5
EXISTING RESET LATCH? P5 OVERHEAD
FAULTS? CSEU +/- 5V DC
CSEU +/- 5V DC
STAB TRIM 28V DC
MENU
ON STAB TRIM 28V DC
OFF
ON
MENU OFF - ALT STAB SW FAULT
- ALT STAB
- CHNL SW FAULT
PIN FAULT
YES NO
YES NO
- CHNL
- 15V PWRPINFAULT
FAULT
- 15VL/R/C
- FCC PWR FAULT
FAULT
- FCC L/R/C FAULT
- L/R/C FSPM FAULT
- L/R/C FSPM FAULT
- YSM FAULT
- YSM FAULT
- AIR/GND
- AIR/GND1/1_2/2
1/1_2/2 FAULT
FAULT TAT
TAT

- L/R ADC FAULT

- L/R ADC FAULT STAB TRIMSTAB TRIM
- YSM WARNING
WARNING
- YSMBUS
BUSXFEED
XFEED r r
- YSM
- YSMVAL
VALDISCRETE
DISCRETE
CAUTION
CAUTION
a a

L/R YSM
L/R YSM

EICAS
EICAS

YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB TRIM CONT.

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YSM - STABILIZER TRIM- FAULT ANNUNCIATION - STAB Air/Ground 1, 1 and 2, 2 Fault

TRIM CONT. The YSM uses majority vote logic to find the in-air or on-ground status of the
airplane. If 2 of 3 air/ground discrete inputs show in air, the YSM uses the in-air
FCC Fault
status. The same condition applies for on-ground. If one air/ground discrete
input is not valid, the YSM uses the last valid value for that discrete and
An FCC fault is set for a FCC bus if a watchdog, parity, coincidence, or engage
continues majority vote. If two air/ground discertes are not valid, the YSM goes
fault is set for that bus. On the ground, FCC faults have an auto reset when data
to in-air status. The YSM on-ground monitor compares the majority voted
becomes valid. In the air, FCC faults have an auto reset when the data becomes
status with each of the three air/ground discrete inputs. If there is a
valid and stays valid for 2.25 seconds.
disagreement for two seconds, a fault is set against the invalid air/ground
discrete. This fault can not reset until the not valid discrete agrees with the
Left, Right, and Center FSPM Fault
majority voting.
The FSPM gives flap and stabilizer position data to the YSM.

If the stabilizer position input is not more than 1v dc for 2 seconds, a left or right
FSPM fault is set. The fault has an auto reset if the stabilizer position
comes back in the correct range.

Comparison and range monitor is done on flap position inputs to YSM software
to make sure of the validity of the data.

If the flap position input value is not in the valid range and stays out of that
range for 2 seconds, a left, center, or right FSPM fault is set. If the airplane is on
the ground and the flap position input comes back to a valid range, the fault has
an auto reset.

Comparison monitor is only done in the air. If a flap position disagrees with the
selected position for 15 seconds, a left, center, or right FSPM fault is set.

YSM Fault

A stabilizer trim YSM fault is set if a trim command coincidence fault (trim
commands between the two microprocessors disagree) or trim mode
coincidence fault (the stabilizer trim mode between the two microprocessors
disagree) occurs. On the ground, a stabilizer trim YSM fault has an auto reset
when the coincidence fault goes away. In the air, a stabilizer trim YSM fault has
an auto reset when the coincidence fault goes away for 15 seconds. After the
stabilizer trim YSM fault has an auto reset four times, the next fault is set and
latched until the airplane is on the ground and the coincidence fault goes away.
 - STCM FAULT
- MAN TRIM SW FAULT
- L/C
STCMHYDR DISCRETE
FAULT
- STAB BRKSW
MAN TRIM SWFAULT
FAULT STAB
L/C HYDR
- STAB 28VDISCRETE
FAULT S Q TRIM
- STAB BRK SW FAULT STAB a
- STAB 28V FAULT S RQ PANEL
TRIM
EXISTING RESET LATCH?
a
P5 OVERHEAD
FAULTS? R PANEL
EXISTING RESET LATCH? P5 OVERHEAD
FAULTS? CSEU +/- 5V DC
CSEU +/- 5V DC
STAB TRIM 28V DC
ON
MENU OFF STAB TRIM 28V DC
MENU
ON
OFF
- ALT STAB SW FAULT
YES NO
-- CHNL PIN SW
ALT STAB FAULT
FAULT
YES NO
- 15V PWR FAULT
CHNL PIN FAULT
- FCC L/R/CFAULT
15V PWR FAULT
- L/R/C FSPM
FCC L/R/C FAULT
FAULT
-- YSM
L/R/CFAULT
FSPM FAULT
YSM FAULT
- AIR/GND 1/1_2/2 FAULT TAT

-- L/R
AIR/GND 1/1_2/2 FAULT
ADC FAULT STAB TRIM
TAT

-- YSM
L/R ADC
BUSFAULT
XFEED WARNING STAB TRIM
-- YSM
YSM BUS WARNING r
VAL XFEED
DISCRETE r
- YSM VAL DISCRETE
CAUTION
CAUTION a
a

L/RL/R
YSM
YSM

EICAS
EICAS

YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB TRIM CONT.

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YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB

TRIM CONT.

The air/ground faults can reset on the ground with one of these common
monitor reset procedures:

- Set the CSEU +5v dc power to off then to on

- Operate the yaw damper engage switch off and on
- Do a RESET LATCH? from the BCM.

Left or Right ADC Fault

This bus failure can be an input data fault (watchdog, parity, SSM, etc.) or a
comparison fault (difference of individual input parameters).

YSM Bus Crossfeed

This fault shows a problem with the wiring between the two YSMs.

YSM Val Discrete

This fault shows when the YSM-YSM IN 1/2 ARINC cross channel bus shows
the other YSM is valid and the cross channel YSM VALID OTHER discrete
shows the other YSM is not valid. If this condition stays for 5 seconds, a
fault is set.
 - STCM FAULT
- MAN TRIM SW FAULT
- STCM
- L/C HYDRFAULT
DISCRETE
- MANBRK
- STAB TRIMSW
SWFAULT
FAULT
- L/C HYDR DISCRETE
STAB
- STAB 28V FAULT S Q TRIM
- STAB BRK SW FAULT STAB a
- STAB 28V FAULT SR Q TRIM
PANEL a
EXISTING
FAULTS?
RESET LATCH? R PANELOVERHEAD
P5
EXISTING RESET LATCH? P5 OVERHEAD
FAULTS? CSEU +/- 5V DC
CSEU +/- 5V DC
STAB TRIM 28V DC
MENU
ON
OFF
STAB TRIM 28V DC
ON
MENU OFF - ALT STAB SW FAULT
YES NO
- ALT STAB
- CHNL SW FAULT
PIN FAULT
YES NO
- CHNL
- 15V PWRPINFAULT
FAULT
- 15VL/R/C
- FCC PWR FAULT
FAULT
- FCC FSPM
- L/R/C L/R/C FAULT
FAULT
- L/R/C
- YSM FSPM FAULT
FAULT
- YSM FAULT
- AIR/GND 1/1_2/2 FAULT TAT
- AIR/GND
- L/R ADC 1/1_2/2 FAULT
FAULT
TAT
STAB TRIM
- L/R ADC FAULT WARNING STAB TRIM
- YSM
- YSMBUS
BUSXFEED
XFEED WARNING r
- YSM r
- YSM VALDISCRETE
VAL DISCRETE
CAUTION
CAUTION
a a

L/R L/R
YSMYSM

EICAS
EICAS

YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB TRIM CONT.

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YSM - STABILIZER TRIM - FAULT ANNUNCIATION -

UNSCHEDULED TRIM
General

One or more of these conditions cause the amber UNSCHED STAB TRIM light
on the P5 overhead panel to come on and a level B caution on the EICAS
display to show:

- Movement of the stabilizer without a valid stabilizer trim command

- Movement opposite to the stabilizer trim command
- Alternate electric trim used when an autopilot is in command.

The annunciation stops when stabilizer motion stops or when the alternate
electric trim stops.

Autotrim Mode

For the autotrim mode, at least one flight control computer (FCC) must be
engaged and no channel code faults or stabilizer trim function faults.

A channel code fault is set if there is a disagreement between the left and right
YSM channel discretes.

A stabilizer trim function fault is set if one of these conditions occur:

- Trim command coincidence fault - stabilizer trim commands from the control
and arm processor channels disagree
- Trim mode coincidence fault - the stabilizer trim active mode from the control
and arm processor channels disagree.

During single or triple channel autopilot operation, the autotrim mode in the left
YSM does a check to make sure the right YSM autotrim mode is disabled.
During dual channel approaches, full rate autotrim is possible.

Unscheduled Stab Trim

If there is stabilizer trim movement without a valid command or opposite to a

valid command, and this condition stays for 0.6 seconds, an unscheduled trim
fault is set and latched. The unscheduled trim fault resets when the stabilizer
movement stops for 3 seconds.
 STAB MOVES
>0.4 DEGREES WITHOUT
VALID
TRIM CMD
0.6 SEC S Q UNSCHEDULED
STAB
MOVES STAB TRIM
STAB 3.0 SEC R UNSCHED
OPPOSITE MOTION STAB TRIM
TO CMD STOPS a
DIRECTION ALTERNATE
ELECT P5 OVERHEAD
TRIM PANEL

STAB TRIM
FUNCTION FAULT AUTOTRIM
CHANNEL CODE MODE
FAULT
ANY FCC ENGA CONTROL PROC
ARM PROC

LEFT YSM TAT

AUTOTRIM WARNING
MODE (RIGHT) r UNSCHD STAB TRIM

CAUTION
a

INTERNALS SAME AS LEFT YSM

EICAS

RIGHT YSM

YSM - STABILIZER TRIM - FAULT ANNUNCIATION - UNSCHEDULED TRIM

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FLT CONT/HYD - GENERAL DESCRIPTION - FLIGHT

CONTROLS

General

All flight control surfaces are moved by hydraulic power control actuators
(PCAs). PCAs are controlled by mechanical linkage with manual, autopilot, and
attitude inputs.

Manual Controls

A system of cables, quadrants, and rods provide manual control from the control
wheel, control column, and rudder pedals. Manual control of the ailerons is
helped by lateral central control actuators (LCCAs) which drive the mechanical
systems in the wings. Manual inputs include artificial feel components to provide
feel in the absence of aerodynamic feedback force.

Autopilot Controls

The autopilot system provides electrical commands to control servos which for
ailerons are the LCCAs. The servos provide the mechanical control of the
PCAs. The required position feedback information is by electrical signal from
the servos.

Attitude Controls

Automatic attitude control is provided for yaw damping. Electrical commands to

the servos result in mechanical motion of the PCA control linkage. Position
feedback is also by electrical signal from the servos.
 LEFT HYD SYS - FLT CONT CENTER HYD SYS - FLT CONT RIGHT HYD SYS - FLT CONT

EDP AC AC AC AC
MP ADP EDP MP
MP MP

L E SLATS
RAT
T E FLAPS

STAB TRIM STAB TRIM

L WING SOV L TAIL SOV C WING SOV C TAIL SOV R WING SOV R TAIL SOV

LCCA LCCA LCCA

SPOILER PCA SPOILER PCA SPOILER PCA
(1-6-12) (3-4-5-8-9-10) (2-7-11)
AIL PCA AIL PCA AIL PCA
(LOB-LIB-ROB) 1 (LIB-RIB) 1 (LOB-RIB-ROB) 1
ELEV A/P ELEV A/P ELEV A/P
SERVO SERVO SERVO
ELEV FEEL ELEV FEEL
CMPTR CMPTR
ELEV PCA ELEV PCA ELEV PCA
(LEFT-RIGHT) (LEFT-RIGHT) (LEFT-RIGHT)
DIR A/P DIR A/P DIR A/P
SERVO SERVO SERVO

RUDDER PCA RUDDER PCA RUDDER PCA

RIGHT YAW LEFT YAW
DAMPER SERVO DAMPER SERVO 1 - LOB - LEFT OUTBOARD
- LIB - LEFT INBOARD
RUDDER RATIO - RIB - RIGHT INBOARD
CHANGER - ROB - RIGHT OUTBOARD

FLT CONT/HYD - GENERAL DESCRIPTION - FLIGHT CONTROLS

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AUTOPILOT CONTROL SERVO This causes the detent pistons to clamp the actuator piston to
the crank.
Purpose
Detent Trip Monitoring
The A/P control servo is a hydraulically powered device which converts an
During the synchronizing portion of engagement and subsequently, the FCC
electrical
compares the autopilot LVDT position with the output LVDT position. If these
signal or command into an output motion or force. This output motion is
two do not agree, this is a condition known as Camout. Camout can occur due
applied to the aileron power control actuators (PCAs) for aileron control. It also
to manual overriding of the autopilot, due to other autopilot control forces during
back drives the manual input to move the control wheel.
multichannel operation, because of mechanical jamming or due to a servo
malfunction.
Servo Actuator Description

The servo consists of two solenoid valves, an electro-hydraulic servo valve

(EHSV), a pressure regulator, hydraulic connections, filter, relief valve, detent
and servo pistons, output crank and autopilot and output LVDT's. The solenoid
valves provide hydraulic pressure to the detent or engage piston and to the
EHSV for control of the servo piston. The pressure regulator reduces system
pressure to the detent piston to allow manual or autopilot override. The EHSV
supplies pressure to one side or the other (C1 or C2) of the servo piston to
cause movement in response to autopilot commands. This movement is
transmitted
via the autopilot LVDT to the flight control computer (FCC). With the
servo fully engaged, the detent pistons are clamped by regulated pressure to
the output crank. Output movement is transmitted to the FCC from the output
LVDT. The autopilot servo piston movement based on autopilot command will
then move the output crank and the aileron PCA's.

Initial Conditions

Before autopilot engagement (as shown), SV1 and SV2 are open and the
detent pistons are retracted by the disengage springs so the output crank is not
clamped to the servo actuator.

Engaged Conditions

When the autopilot is first engaged, the FCC energizes SV1. This supplies
hydraulic pressure to the EHSV and to SV2. With the EHSV enabled,
commands from the FCC synchronize the actuator piston with the output crank.
This is accomplished by using inputs from the autopilot LVDT and output LVDT.
With synchronizing complete and other engage requirements satisfied, the FCC
energizes SV2.
 SV 2 ELECTRICAL
CONNECTOR
EHSV
SV 1
FILTER AUTOPILOT LVDT
COVER

RETURN
PRESSURE
SOLENOID
OUTPUT SOLENOID ELECTRO-HYDRAULIC VALVE 2
CRANK VALVE 1 SERVO VALVE (EHSV) ELECTRICAL
HYDRAULIC CONNECTOR
RETURN PORT
PRESSURE
PR C P C1 C2 R PRC REGULATOR
OUTPUT LVDT
FILTER
P
R
HYDRAULIC C
PRESSURE PORT ROLLER RP RELIEF
ORIFICE (2) VALVE

SERVO PISTON
CENTERING SPRING
DETENT ACTUATOR
SERVO PISTON PISTON LVDT
DISENGAGE
OUTPUT LVDT SPRINGS
DETENT INTERNAL
PISTON CRANK
PIVOT FOR OUTPUT
CRANK & INTERNAL OUTPUT CRANK
CRANK TO AILERON
CONTROL LINKAGE

AUTOPILOT CONTROL SERVO

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LATERAL CENTRAL CONTROL ACTUATORS The second enable signal opens solenoid valve 2 and hydraulic pressure goes
to the detent piston. The detent piston then closes onto the roller on the input
crank and further movement of the autopilot actuator piston results in movement
General Description
of the booster crank.
The lateral central control actuator (LCCA) has a booster section and an
Servo Details
autopilot section. The booster section is controlled by the input arm during
control wheel operation or by the autopilot section when the FCC is engaged.
The control actuator pistons are centered and held open by spring pressure
The booster then moves the wing cables.
when the autopilot is not engaged.
The detent piston is closed by regulated pressure which is equalized between
LCCA Operation - Control Wheel
the two sides.
The autopilot actuator piston is moved by differential pressure between the two
The movement of the booster section piston moves the wing cables. The smaller
sides as provided from the EHSV.
side of the unbalanced area piston receives a constant hydraulic pressure of
The 26v ac excitation of each LVDT is provided from the associated flight
1500 psi from a pressure regulator and relief valve. The control valve also
control computer.
receives 1500 psi from the pressure regulator and relief valve, and ports
metered pressure to the larger side of the unbalanced area piston. As the
Autopilot LCCA Operation
primary cables move, the input crank rotates about point A. This rotates the
summing lever about point D and moves the control valve. Movement of the
Before the autopilot is engaged, the LCCA operates as a power booster for
piston changes the position of point D which rotates the summing lever about
manual control and the control section is electrically and hydraulically inactive.
point C and nulls out the input to the control valve.
When the CMD push button on the MCP is pushed, the flight control computer
(FCC) supplies an aileron hydraulic arm discrete (dc) to servo valve 1 (SV1).
A relief valve and a by-pass check valve allow backdrive of the piston by the
The FCC provides commands to the EHSV to synchronize the autopilot actuator
other LCCAs. An anti-jam detent allows the input crank to break out (10 lbs
piston (actuator LVDT) with the booster actuator piston (output LVDT).
control wheel) if the control valve jams and allows operation of the other LCCAs.
When synchronized (1 second) and when other engagement requirements are
Hydraulic pressure to the booster section is available when its hydraulic system
complete, the FCC provides another dc discrete to engage the detent pistons.
is pressurized. No electrical signal is necessary.
The autopilot actuator piston is then clamped to the input crank roller (Point B).
LCCA Operation - Autopilot
LCCA Operation
The autopilot portion of the servo responds to autopilot commands. The servo is
Autopilot commands for aileron deflection are received by the EHSV which
enabled by two solenoid valves and controlled by an electrohydraulic servo
moves the autopilot actuator piston. This movement is translated through the
valve (EHSV). Detent pistons engage the actuator piston to provide servo
input booster crank to both the input arm and the crank. Movement of the
output. The actuator piston position is detected by the autopilot LVDT.
booster crank moves the booster actuator and moves of the aileron surface.
Hydraulic pressure on the detent pistons is decreased by the pressure regulator
Servo Enabling
to allow manual control inputs to override autopilot inputs. Manual force can
move the detent pistons in the autopilot actuator piston and the booster crank
Two discrete signals are necessary to enable the servo. The first opens
can be moved. This causes LVDT disagreement and a detent trip which is
solenoid valve 1 and provides hydraulic pressure to solenoid valve 2 and the
detected by the FCC.
EHSV.The EHSV then controls the position of the autopilot actuator piston.
When the actuator LVDT position agrees with the output LVDT position, the flight
control computer can provide the second enable signal.
 LATERAL CENTRAL CONTROL ACTUATORS
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AFDS GENERAL - AFDS - GENERAL DESCRIPTION Major Components

General The AFDS has these components:

The autopilot flight director system (AFDS) gives flight director commands or - Autoflight control systems mode control panel (AFCS MCP)
automatic control of the airplane in all phases of flight. The AFDS makes - Flight control computers (FCCs)
stabilizer trim commands when an autopilot is in command. - Caution and warning annunciators
- Autoland status annunciators
Functional Operation - Autopilot servos.

These are the three main tasks done by the AFDS:

- Pilot assist modes

- Pilot-controlled command modes
- Automatic modes.

The pilot assist modes are the flight director modes.

These are the pilot-controlled command modes:

- Vertical speed
- Altitude hold
- Heading select
- Heading hold.

These are the automatic modes:

- Coupled approach
- Autoland
- Rollout
- Go-around
- Vertical and lateral navigation with the flight management system
- Automatic stabilizer trim with the stabilizer trim system.
 AFDS GENERAL - AFDS - GENERAL DESCRIPTION
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AUTOPILOT FLIGHT DIRECTOR SYSTEM

General

The autopilot and flight director system (AFDS) provides automatic guidance
and control of the airplane ailerons, elevators, and, for autoland, the rudder. It
also provides the guidance processing for the flight director displays.
Engagement and mode selection of the autopilot and flight director are
accomplished through the AFCS mode control panel.

Annunciation and Display

Each attitude director indicator (ADI) portion of the PFD displays flight director
and AFDS commands on the flight mode annunciator (FMA). The FMA
identifies AFDS status, roll/pitch arm and engaged modes. The autoland status
annunciators identify the system capability and limitation status for autoland
operations. The AFCS mode control panel has mode selector switch
annunciation and reference readouts. A red A/P DISC light and an amber
AUTOPILOT caution light gives visual alerts for A/P warnings and cautions.
 A/T ARM A/P ENGAGE
IAS/MACH HDG VERT SPD ALT L C R

F/D L NAV B/CRS CMD CMD CMD F/D

IAS
ON OFF ON

SEL 5 25
THR V NAV
BANK DN LOC
OFF AUTO SEL OFF
LIMIT

SPD FL CH HOLD V/S HOLD APP DISENGAGE

UP

AFCS MODE CONTROL PANEL

EADI (2)

ELECTRONIC FLIGHT INSTRUMENT

SYSTEM (EFIS) EHSI (2)

AUTOLAND STATUS
ANNUNCIATOR (2)
AUTOPILOT DISENGAGE A/P DISCONNECT LIGHT
SWITCH (2)

A/P CAUTION LIGHT

GO-AROUND SWITCH (2) A/P ROLLOUT GUIDANCE

SERVO (3)

A/P LATERAL CONTROL

SERVO (3)
A/P PITCH CONTROL
ELECTRICAL POWER SERVO (3)
FLIGHT MANAGEMENT SYSTEM
AIR DATA/INERTIAL REFERENCE SYSTEM
RADIO ALTIMETER YSM (2)
INSTRUMENT LANDING SYSTEM
CENTRAL WARNING SYSTEM
ENG IND CREW ALERTING SYSTEM
FLIGHT CONTROL MCDP
COMPUTER (3)

AUTOPILOT FLIGHT DIRECTOR SYSTEM

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AFDS GENERAL - BLOCK DIAGRAM Flight Controls Interface

General The autopilot servos give the mechanical power to operate the surface power
control actuators (PCAs). The autopilot gives stabilizer trim with the yaw
This block diagram shows the AFDS hardware configuration by function. damper/stabilizer trim modules (YSMs) and the stabilizer trim control modules
(STCMs).
Pilot Control Function
Electronic Flight Instrument System
The AFCS mode control panel gives the primary interface between the pilots
and the AFDS. The thrust management system also uses the mode control The EADI is the primary display for the AFDS and shows these functions:
panel. Pilot inputs have a direct interface with the flight control computers.
- Flight director commands
I/O, Calculation, and Monitor Functions - Flight mode annunciation
- Status.
The flight control computers (FCCs) are the primary calculation units of the
AFDS. These are the functions of the FCCs: The EHSI shows the selected heading index.

- Control law calculation Maintenance Function

- Sensor interface and management
- Control system interface The maintenance control and display panel (MCDP) gives the maintenance
- Engage and disengage function for the AFDS.
- Monitor
- Annunciation management.

Navigation Sensors

The avionics navigation sensors supply inertial, atmospheric, and ground

reference feedback data to the AFDS. The navigation sensors necessary for
autoland are triple redundant. The other navigation sensors are dual
redundant.

Airplane Configuration Sensors

The airplane configuration sensors give the necessary data to make

modifications in the control in relation to the airplane flight properties.
 TMC DISCRETE WARNING
INDICATORS
AFCS MODE CONTROL PANEL ASA
DISCRETE CAUTION
EADI INDICATORS
EHSI
STAB POSN FMC
EFIS TO
DUAL REDUNDANT ELEV SURFACE
SERVOS PCA(S)
NAVIGATION SENSORS
FLAP POSN TO
AILERON SURFACE
SERVOS PCA(S)
TO
SPD BRAKE DIRECT SURFACE
HDLE POSN SERVOS PCA(S)
FLIGHT
CONTROL AUTOPILOT SERVOS
COMPUTER
SLAT SW
LOGIC ADIRU
YSM STCM

AUTO STAB
TRIM
MCDP
HYD VALID ILS
LOGIC
DISENGAGE SWITCHES MAINTENANCE
FUNCTION
DFDAU
AIR/GND F/D
RA
LOGIC SOURCE HDG REF FLIGHT
SEL SW SWITCH RECORDER
AIRPLANE CONFIG TRIPLE REDUNDANT G/A SWITCHES FUNCTION
SENSORS NAVIGATION SENSORS MISCELLANEOUS PILOT INPUTS

AFDS GENERAL - BLOCK DIAGRAM

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AFDS GENERAL - COMPONENT LOCATION

Flight Compartment Components

The flight compartment components for AFDS operation give these functions:

- Control
- Flight director display and source selection
- Annunciation
- Warning and caution alerts
- Autopilot disconnect
- Go-around functions.

Airborne Data Loader

The airborne data loader (ADL) is a general purpose disk drive unit that does
uploads and downloads of computer programs and data. The ADL connects to
the airplane computers by ARINC 429 data buses. Connection to LRUs is at the
data loader control panel.

Main Equipment Center Components

The FCCs are in the main equipment center.

Autopilot Servos

The three lateral central control actuators are in the wing roots with the left and
right units on the left side and the center unit on the right side.

The three elevator autopilot servos are in the tail aft of the stabilizer.

The three directional autopilot servos are in the vertical stabilizer.

 P1-3 PANEL FLAP LEVER
- A/P DISC LIGHT - LEADING EDGE
- A/P CAUTION SLAT SWITCH
LIGHT
P55 GLARESHIELD PANEL
- MODE CONTROL PANEL
THRUST LEVERS
- GO-AROUND P3 F/O INSTRUMENT PANEL
SWITCHES - EADI
- INSTRUMENT SOURCE
SELECT SWITCH PANEL
- ASA
- HEADING REFERENCE SWITCH
P61 AUXILIARY PANEL
- AIRBORNE DATA LOADER

YAW DAMPER SERVOS

P1-1 CAPT INSTRUMENT PANEL CONTROL WHEELS
- EADI - A/P DISCONNECT
- INSTRUMENT SOURCE SWITCHES ELEV A/P SERVOS DIRECTIONAL
SELECT SWITCH A/P SERVOS
- ASA

L FCC (E1-3) LATERAL CENTRAL

CONTROL ACTUATORS
C FCC (E1-4)
R FCC (E1-5)

FWD

AFDS GENERAL - COMPONENT LOCATION

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MODE CONTROL PANEL - FUNCTIONS HEADING DISPLAY- This HEADING DISPLAY is capable of displaying 000 to
359 degrees in increments of one degree. Initialized at 000 on power up.
Switches, Lights and Controls
HEADING Select KNOB- The outer knob has 6 positions for bank angle limits.
The inner knob controls the heading of aircraft and heading bug on EFIS.
Autothrottle Arm Sw- Autothrottle is Armed by use of this toggle switch when in
Depressing SEL causes the aircraft to obtain the heading selected on the
A/T ARM position. This switch provides DC PWR to the autothrottle servo. OFF
window. Another way is to press select then turn the heading to the desired
disables autothrottle Autothrottle Mode Select switches. The FMA displays
heading and the A/P will follow as the knob is turned.
autothrottle mode of operation.
HEADING HOLD SWITCH- This flowbar illuminated indicates HDG has been
The following are Autothrottle modes:
held.
THR -When the THR light bar is on engine(s) are at MAX EPR or MAX Thrust
VERTIICAL SPEED DISPLAY- The VERTIICAL SPEED DISPLAY window
as selected from the TMSP.
displays Vertical Speed (V/S) from +6000 to -8000 ft/min in 100 ft/min
increments.
SPD- When the SPD light bar is on engine(s) are maintaining the aircraft speed
selected from the IAS/MACH speed window. Limits are selected from the TMSP.
VERTICAL SPEED RATE WHEEL-dials in a selected V/S.
VNAV- When the VNAV light bar is on engine(s) Limits, Speed, throttle position,
VETTICAL SPEED SLECET SWITCH-The V/S switch activates the V/S mode.
pitch of the aircraft and performance are being determined by the FMC.
ALTITUDE DISPLAY WINDOW- ALT shows from 0-50,000 feet in 100 feet
increments.
FL CH- When the FL CH light bar is on engine(s) are maintaining the aircraft
speed selected from the SPEED DISPLAY window. Limits are selected from the
ALTITUDE SELECT KNOB/SWITCH-Turning of the knob clockwise increases
TMSP.
the ALT in the ALTITUDE DISPLAY WINDOW.
Lateral Navigation-LNAV light bar is when LNAV is either ARMED or Engaged
ALTITUDE HOLD-HOLD This flowbar when pressed engages ALT hold.
as displayed on the FMA. When L NAV is operating the aircraft roll and heading
is being controlled by the FMC.
AUTOPILOT DISENGAGE BAR-DISENGAGE is a switch that removes power
from the AP servos by applying a ground to the servos.
SPEED DISPLAY - This speed displays 100-399 knots in 1 knot increments and
from .40 to .95 Mach in 0.01 increments. Used in Speed and FL CH mode.
AUTOPILOT ENGAGE SWITCHES- This CMD flowbar illuminated indicates
that the L, R and/or C Autopilot (AP) is engaged or in an armed condition. If
AIRSPEED / MACH CONTROL Switch- changes the speed display numbers
APP has been selected it is possible to have multiple CMD flow bars
and transfers the command to and from the FMC to the MCP when in
illuminated. These AP(s) do not engage until less than 1500 feet RA.
performing a speed intervention.
BACK COURSE SELECT- This flowbar illuminated indicates Back course
AIRSPEED / MACH SEL- SEL changes the display between airspeed and
selected
mach. Mach must be above .40 to display mach.
LOCALIZER SELECT- This flowbar illuminated indicates LOC selected.
Flight Director switches- F/D controls the FD displays on the left and right PFD
ADI respectively.
APPROACH SELECT- This flowbar illuminated indicates APP selected.
 VERTICAL SPEED BACK COURSE SELECT SWITCH/LIGHT
SPEED DISPLAY HEADING DISPLAY DISPLAY
MACH / IAS LATERAL
AUTOTHROTTLE NAVIGATION
ARM SWITCH MODE SELECT HEADING SELECT AUTOPILOT ENGAGE
SWITCH/LIGHT ANK ANGLE ALTITUDE DISPLAY SWITCHES/LIGHTS
LIMIT SELECTOR

THRUST MODE
SELECT SWITCH FLIGHT VERTICAL SPEED
LEVEL CHANGE RATE UP/DOWN LOCALIZER SELECT FLIGHT DIRECTOR
MODE SELECT SELECT WHEEL SWITCH/LIGHT ON/OFF SWITCH (2)
VERTICAL
SWITCH/LIGHT NAVIGATION
SPEED MODE APPROACH SELECT
SELECT SWITCH MODE SELECT VERTICAL SPEED SELECT
SWITCH/LIGHT SWITCH/LIGHT
SWITCH/LIGHT AUTOPILOT
AIRSPEED/MACH DISENGAGE
CONTROL AND ALTITUDE SELECT SWITCH/LEVER
LOCAL/REMOTE SPEED SWITCH
AIRSPEED / MACH SELECT SWITCH HEADING HOLD SELECT
SELECT SWITCH SWITCH/LIGHT
ALTITUDE HOLD
SWITCH/LIGHT

MODE CONTROL PANEL - FUNCTIONS

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AFDS - ANNUNCIATION AND WARNING - GENERAL

DESCRIPTION
Description

All three FCCs send this data to each EFIS symbol generator:

- Mode status
- Flight director commands
- Heading cursor.

Mode status and flight director commands show on the EADI. The heading
cursor shows on the EHSI. The EFIS symbol generators control the data source
selection. Autoland status LAND 3 or LAND 2 shows in green on the
Autoland Status Annunciator or ASA.

The autopilot disengage warning light signal from each flight control computer
goes to the panel light and the warning electronics unit. The warning electronics
unit causes the captain and first officer master warning lights to come on and
supplies the two-tone siren aural warning.

The autopilot caution light signal from each flight control computer goes to the
panel light and to each EICAS computer. The EICAS computer causes the
captain and first officer master caution lights to come on and
gives a signal to the warning electronics unit. The warning electronics unit
supplies the level B caution aural.

Autopilot disengage and autopilot caution message signals from each flight
control computer goes to each EICAS computer. The EICAS computer shows
the necessary level A or level B message on the applicable display unit.

Each flight control computer supplies autopilot multichannel operation status

signals during an approach mode to the captain and first officer autoland status
annunciators. The autoland status annunciators give four messages for the
different multi-channel operation conditions.
 A/T ARM A/P ENGAGE
WARNING IAS/MACH HDG VERT SPD ALT L C R

F/D L NAV B/CRS CMD CMD CMD F/D

IAS
ON OFF MACH ON
CAUTION
SEL 5 25
THR V NAV
LOC
CAPTAIN AND F/O OFF AUTO SEL BANK
LIMIT
DN
OFF

MASTER WARNING SPD FL CH HOLD V/S HOLD APP DISENGAGE

AND CAUTION UP
LIGHTS (P7)
MODE CONTROL PANEL (P55)
STATUS MODE ANNUNCIATION (TYP)
PITCH MODES G/S LOC CMD
OPERATE FLARE ROLLOUT CMD AUTOLAND STATUS
1
F/D NO LAND 3 (AMBER) LAND 2 (GREEN)
P/RST
TO NO AUTOLAND (AMBER) TEST LAND 3 (GREEN)
EADI TST 2
ALT HOLD
PITCH MODES ROLL MODES
ALT CAPT ARMED ARMED OPERATE AUTOLAND STATUS
AUTOPILOT ANNUNCIATION (P1-3)
V/S G/S LOC TO WARNING RED
SPD FLARE ROLLOUT ANNUNCIATION
HDG HOLD (RED)(P1-3) TAT +12C
G/S VNAV B/CRS HDG SEL TO
AUTOPILOT DISC
FLARE AUTOPILOT A/P
LNAV LNAV CAUTION AUTOPILOT
DISC
GA LOC ANNUNICATION
(AMBER)(P1-3)
VNAV SPD TRK 252 M
B/CRS
VNAV PTH ROLLOUT AUTO YELLOW
SPD LIM TUKWA PILOT
ONP ATT
FLAP LIM GA
VNAV ALT EICAS UPPER DISPLAY UNIT
HEADING KEY:
CURSOR BENSN
MODES ENGAGED ARMED LIMIT
STATUS GREEN N/A N/A
20

PITCH GREEN WHITE GREEN

ROLL GREEN WHITE N/A
EHSI
AFDS - ANNUNCIATION AND WARNING - GENERAL DESCRIPTION
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AFDS GENERAL - AUTOLAND STATUS ANNUNCIATORS

General

The autoland status annunciators provide information on degradation from a

triple redundant autoland capability.

Component Details

There is an upper and a lower display on each autoland status annunciator. The
upper display shows the autopilot system autoland status (when in multi-
channel autoland) and the lower display shows the degradation from land 3
capability.

Each display has three faces; blank face, A face, and B face. The displays
operate by magnetic coils. There are two test switches. Push TEST 1 to show
the upper and lower display A faces. Push TEST 2 to show the B faces. The
reset (P/RST) switch is used to clear the lower display (not during test).

EICAS Status Messages

A degraded autoland system failure that occurs below 200 feet shows as a NO
LAND 3 status message after these conditions are true:

- Airplane is on the ground.

- Ground speed is less than 40 knots.
- Autopilot is disconnected.
 LAND 3 RESET SWITCH
(A) TEST SWITCH
(GREEN) (A FACES)
AUTOLAND STATUS AUTOLAND STATUS
1 (B) LAND 2 1
P/RST (GREEN) P/RST
TEST TEST

2 NO LAND 3 2
(A) (AMBER)

(B) NO AUTOLAND TEST SWITCH

(AMBER) (B FACES)

AFDS GENERAL - AUTOLAND STATUS ANNUNCIATORS

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AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY Autopilot Channel Selection

SEQUENCE -1 Autoland status annunciations are different with the
number of autopilot channels engaged. Displays and
General
related operations are described for single, dual, and
triple channel selection.
Each FCC gives its status to the other FCCs on the cross channel data bus.
Use of its status and the status received from other FCCs supplies the
applicable display. Priority programming limits FCC output to one status signal.
Priority sequence is NO AUTOLAND, NO LAND 3, LAND 2, and LAND 3.

Display Groups

ASA displays are divided into two groups:

- Degradation
- Status and degradation.

Degradation displays are provided anytime from airplane on the ground prior to
takeoff until the airplane has descended to 1500 feet of final destination.

Status and degradation displays are provided after multi-channel engaged and
descent below 1500 feet until the airplane lands and the FCCs are disengaged.

Display Sequence

Displays and related functions are divided into these different areas:

- On ground
- In air, climb, or cruise
- Approach select
- Below 1500 feet
- Below 600 feet
- Below alert height
- Land and FCC disengage.

Each area contains displays for single fault and multiple fault conditions.
Automatic operation and manual operation that result from these conditions are
described.
 DEGRADATION DISPLAYS CRUISE AUTOLAND STATUS
1
CLIMB IN AIR, CLIMB, OR CRUISE P/RST
TEST

IN AIR - MONITOR OF FCC, SENSORS AND INTERFACES 2

- NO AUTORESET
ON GROUND FAULT CONDITION
SINGLE FAULT BEFORE APPROACH SELECT
- ON GROUND MONITOR DYNAMIC
TEST OF SENSORS AND INTERFACES

SINGLE FAULT NO LAND 3 NO LAND 3

- MANUAL RESET PUSH AND HOLD - AUTORESET WHEN

REINTIALIZES VOTING CIRCUIT APPROACH SWITCH PUSHED
NO LAND 3 FOR REASSESSMENT OF CONDITION - VOTING CIRCUIT REINITIALIZED
FOR REASSESSMENT OF CONDITION
- AUTORESET WHEN
CONDITION CLEARED

- DISPLAY REMOVED DURING RESET - DISPLAY REMOVED IF FAULT

- DISPLAY REMAINS BLANK WHEN RESET CONDITION CLEARED BEFORE
RELEASED IF FAULT CONDITION CLEARED APPROACH SELECTED
- DISPLAY REMOVED

MULTIPLE FAULTS MULTIPLE FAULTS MULTIPLE FAULTS

NO AUTOLAND NO AUTOLAND NO AUTOLAND

- AUTORESET WHEN - MANUAL RESET - AUTORESET WHEN

CONDITION CLEARED APPROACH SWITCH PUSHED

- DISPLAY REMOVED - OPERATION SAME AS FOR - OPERATION SAME AS FOR

NO LAND 3 ABOVE NO LAND 3 ABOVE
AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE -1
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AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY

SEQUENCE - 2

General

This page continues the autoland status annunciator

display sequence.
 STATUS AND DEGRADATION DISPLAYS APPROACH
AUTOLAND STATUS
SELECT BELOW 1500 FT 1
FAULT CONDITION
AFTER APPROACH SELECT TRIPLE CHANNEL P/RST
TRIPLE TEST
ENGAGE
CHANNEL BELOW 2
SINGLE FAULT ARM 600 FT 1
BELOW
ALERT HEIGHT
NO LAND 3

- MANUAL RESET BELOW LAND AND FCC

BELOW 1500 FT ALERT HEIGHT DISENGAGE
LAND 3 LAND 3

NO LAND 3

- DISPLAY REMOVED AND LATCHED - MANUAL RESET

OFF UNTIL LANDING AND DISENGAGE SINGLE FAULT SINGLE FAULT REVERTS TO ON
- FAULT CONDITION LATCHED IN FCC GROUND MONITOR
- NO UPGRADE FOR CLEARED FAULT LAND 2 LAND 3
CONDITION
NO LAND 3

MULTIPLE FAULTS - MANUAL RESET MULTIPLE FAULTS

NO AUTOLND

LAND 2
- MANUAL RESET
NO AUTOLND NO AUTOLND REVERTS TO ON
GROUND MONITOR
- DISPLAY LATCHED ON UNTIL - NO LAND 3 DISPLAY - NOT REMOVED BY
LANDING AND DISENGAGE REMOVED WITH MANUAL RESET
- DISPLAY REMOVED DURING MANUAL CONDITION EXISTING
RESET WHILE SWITCH IS HELD MULTIPLE FAULTS
- FAULT CONDITION LATCHED IN FCC 1 IF MULTICHANNEL IS NOT ENGAGED BY
- NO UPGRADE FOR CLEARED 600 FT, THE ASA SHOWS NO AUTOLAND.
FAULT CONDITION NO UPGRADE FOR CLEARED FAULT CONDITION.
NO AUTOLND
ASA AUTORESETS WHEN A/P IS DISENGAGED
OR NEW MODE IS SELECTED.
- NOT REMOVED BY
MANUAL RESET

AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE - 2

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AFDC - FUNCTIONAL DESCRIPTION - ASA - DISPLAY

SEQUENCE - 3
General

This page continues the autoland status annunciator

diaplay sequence.
 CRUISE
APPROACH AUTOLAND STATUS
SELECT AND 1
DUAL CHANNEL ARM P/RST
TEST
APPROACH SELECT - DUAL CHANNEL ARM BELOW 1500 FT AND
- MONITOR OF FCC, SENSORS AND INTERFACES DUAL CHANNEL ENGAGE 2

SINGLE FAULT CONDITION BELOW 600 FT

STATUS AND
AFTER APPROACH SELECT BELOW DEGRADATION DISPLAYS
AND DUAL CHANNEL ARM ALERT HEIGHT
SINGLE FAULT
BELOW 1500 FT AND BELOW LAND AND FCC
DUAL CHANNEL ENGAGE ALERT HEIGHT DISENGAGE
NO LAND 3 LAND 2 LAND 2

NO LAND 3 NO LAND 3 NO LAND 3

- ASA REMAINS BLANK
- FAULT CONDITION LATCH IN FCC(S) SINGLE FAULT SINGLE FAULT - RESET REVERTS TO
MULTIPLE FAULTS ON GROUND MONITOR

LAND 2 LAND 2

NO AUTOLND NO LAND 3 NO LAND 3 NO AUTOLND

- DISPLAY LATCHED ON UNTIL - FAULT CONDITION - FAULT CONDITION - RESET REVERTS TO

LANDING AND DISENGAGE LATCH IN FCC(S) 1 LATCH IN FCC(S) 1 ON GROUND MONITOR
- DISPLAY REMOVED DURING MANUAL MULTIPLE FAULTS MULTIPLE FAULTS
RESET WHILE SWITCH IS HELD
- FAULT CONDITION LATCHED IN FCC(S)
- NO UPGRADING FOR CLEARED
NO AUTOLND NO AUTOLAND
FAULT CONDITION
- NOT REMOVED BY - NOT REMOVED BY
MANUAL RESET MANUAL RESET
- IMMEDIATE NO
AUTOLAND
1 NO LAND 3 SHOWS AFTER
LANDING AND DISENGAGE.

AFDC - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE - 3

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AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY

SEQUENCE - 4
General

This page continues the autoland status annunciator display sequence.

 STATUS AND DEGRADATION DISPLAYS
SINGLE CHANNEL OR FLIGHT DIRECTOR
BELOW
ALERT HEIGHT

BELOW LAND AND FCC

ALERT HEIGHT DISENGAGE

NO LAND 3

SINGLE FAULT - RESET REVERTS TO

ON GROUND MONITOR

NO AUTOLND

- FAULT CONDITION - RESET REVERTS TO

LATCHED IN FCC(S) ON GROUND MONITOR
1
MULTIPLE FAULTS

NO AUTOLND

- NOT REMOVED BY
MANUAL RESET

NO LAND 3 SHOWS AFTER

1
LANDING AND DISENGAGE
AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE - 4
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MODE CONTROL PANEL INTERNALS Monitors

Power supply monitoring hardware supplies appropriate reset signals to the

General
CPU when power supplies have fallen below minimum values. A compute-cycle
monitor verifies that the CPU is capable of periodic execution to re-start a
To accomplish the requirement that a single failure will not cause total loss of
hardware watchdog timer. If the timer is not restarted every 150 to 200
the AFDS mode control panel function, the mode control panel contains two
milliseconds, the critical portions of the I/O system and associated CPU are
separate control channels with a common display. Each channel is powered
disabled allowing the alternate CPU to assume control without interference.
from a different airplane 28 volt dc source.

The A channel is supplied by the left dc bus and the B channel is supplied by
the right dc bus. The A channel receives ARINC 429 data from the left and
center flight control computers (FCC) and from the thrust management
computer (TMC).

The B channel receives ARINC 429 data from the right and center flight control
computers and from the thrust management computer.
Each channel independently processes the 429 bus received data. The pilot
input front panel data drives its own ARINC 429 bus which transmits that data to
the appropriate receiving computers.

Front Panel Switches

Each channel receives data from each switch or encoder on the front panel.
Electrical isolation is maintained between each switch contact set and each
encoder's A and B channel outputs to ensure system redundancy.
The autopilot disengage bar and the autothrottle arm switch provide a positive
disengagement function.

Input and Output Discretes

A master test input to each channel allows for external testing of the lights for all
dot bars simultaneously.

Front Panel Displays

Each channel can control the common numerical displays since each channel's
bi-directional data buffers are connected together. Only a single channel will
provide data to the displays; the channel not in command is in the read mode
(high impedance) to avoid potential conflicts.
 MODE CONTROL PANEL INTERNALS
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ENGAGE LOGIC - SIMPLIFIED Elevator and Aileron Engage

The engage function is initiated when the arm state is achieved. The elevator/
General
aileron engage relay operates similarly to the arm relay to energize the engage
solenoid in the elevator and lateral autopilot servos.
Engagement of the autopilot servos is a two-step process. Arm status is
achieved first if all arm conditions are met and removal conditions are not
Rudder Arm
present. Engage status is achieved second if engage removal conditions or
faults do not exist. Details of the logic statements shown here are developed
later. The engagement process is comprised of two operations elevator/aileron Rudder arm is initiated by multichannel operation which is enabled only in the
engagement and rudder engagement. approach mode. The rudder arm solenoid is energized through the rudder arm
relay similar to elevator/aileron arm.
Elevator/Aileron Engagement
Rudder Engage
Arm is initiated by command request from the MCP if single channel engaged
and by command request and approach mode with another channel engaged. Upon achieving rudder arm status, the rudder engage logic is initiated leading to
With the arm latch set, the elevator/aileron arm relay is energized which will the energizing of the rudder engage solenoid when multichannel engaged.
energize the arm solenoid. The arm latch also enables the engage logic. If
conditions are valid for engagement, the engage latch will be set energizing the
elevator/aileron engage relay. When the engage latch is set, the arm latch is
bypassed and arm removal faults will not disengage the autopilot.

Rudder Engagement

Rudder engagement is initiated by multi-channel arm during approach. This

sets the rudder arm latch and, like the elevator/aileron engagement, the arm
relay and arm solenoid are energized and the engage logic is enabled. Rudder
engage is achieved when two channels are engaged if local channel is one of
them. Notice that rudder engagement is removed if local channel disengages or
if there is no foreign channel engaged.

ENGAGE INTERLOCKS - OPERATION

Elevator and Aileron Arm

The arm function is initiated by pressing the CMD pushbutton. This causes a
request to be transmitted via the data bus to the arm logic. When the arm state
is achieved, a solid state switch is closed, energizing the elevator/aileron arm
relay. This relay supplies 28v dc to energize the elevator and aileron arm
solenoids within the autopilot servos. The arm state is also transmitted to the
MCP to illuminate the CMD pushbutton.
 28V DC
SERVO ELEV ARM ELEVATOR ARM
SOLENOID
AIL ARM
ARM & ENGA DISC
ELEV/AIL ELEVATOR
ARM RLY ENGAGE
SOLENOID
429 429 ELEV & AIL ELEV
XMTR RCVR ARM LOGIC SERVO
DATA ENGAGE A/P PITCH
MGMT CONTROL SERVO
429 429
RCVR XMTR
D AILERON ARM
AIL SOLENOID
A ELEV/AIL SERVO
T ENGA RLY ENGAGE
A ELEV & AIL
CMD AILERON
M ENGAGE ENGAGE
A LOGIC SOLENOID
N RUDDER
A ARM
G RUDDER A/P LATERAL
E ARM RLY CONTROL SERVO
OFF M
E RUDDER ARM
DISENG BAR LOGIC
N RUDDER ARM
AFCS MCP T RUDDER SOLENOID
SERVO
ENGAGE
RUDDER RUDDER
ENGA RLY ENGAGE
X-CHANNEL 429 RUDDER SOLENOID
BUSSES RCVR ENGAGE
LOGIC

A/P ROLLOUT
FLIGHT CONTROL COMPUTER (TYP) GUIDANCE SERVO

ENGAGE LOGIC - SIMPLIFIED

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ENGAGE LOGIC DETAILS - HARDWARE MONITORS If the autopilot is dual channel engaged, a camout will cause an autopilot
disconnect annunciation with a no autoland annunciation on the ASA.
If the autopilot is triple channel engaged, a single channel camout will cause; if
General
above the alert height, a LAND 2/NO LAND 3 annunciation on the ASA; if below
the alert height, the ASA will annunciate LAND 3 and there will be no changes
The hardware monitor logic receives the arm or engage logic developed
until touchdown and disconnect. The ASA will then annunciate NO LAND 3.
previously and if conditions are appropriate, passes this logic to the arm and
If the autopilot is triple channel engaged a dual or triple channel camout will
engage and for rudder arm and engage.
cause an autopilot disconnect annunciation and NO AUTOLAND on the ASA.
Elevator/Aileron Arm

For elevator/aileron arm, wheel disengage, computer valid and power supply
valid are monitored.

Elevator/Aileron Engage

For elevator/aileron engagement, the same things are monitored. In addition,

camout (local and foreign) is also monitored. Camout is determined within the
local channel by the detent comparator. Detent logic remains high as long as
the servo LVDT and surface LVDT agree within the limits shown. MY DETENT
logic low will cause a camout disengage only if there is another channel servo
engaged.

Relative left or right channel camout (detent low) will cause disengagement if
the foreign channel disengages with a detent trip and it is the only other channel
engaged. In other words, for dual configuration if one channel has a detent trip,
both channels disengage. Rudder Arm For rudder engagement, ELEV/AIL
PREARM and power supply valid is necessary in addition to ROLLOUT ARM.

Rudder Engage

For rudder engagement logic, ELEV/AIL PRE-ENGA and FOREIGN SERVO

ENGA are necessary in addition to ROLLOUT ENGA and power supply valid.

AFDS Annunciations

If the autopilot is single channel engaged, a camout will cause an autopilot

caution annunciation.
 DC ELEV/AIL
WHEEL DISENGAGE POWER ARM RELAY
(FROM ELEV/AIL ARM LOGIC) ELEV/AIL ARM SUPPLY VALID

COMPUTER VALID PRE ARM

FROM COMPUTER
MONITOR COMPUTER INHIBIT 1
A
POWER
WHEEL DISENGAGE SUPPLY VALID DC ELEV/AIL
(FROM ELEV/AIL ENGA LOGIC) ELEV/AIL ENGAGE ENGA RELAY
LEFT DETENT LEFT CAMOUT PRE ENGA
FROM OTHER LEFT SERVO ENGA CAMOUT
CHANNELS LOGIC
RIGHT DETENT RIGHT CAMOUT B
RIGHT SERVO ENGA 2
AIL SERVO
AIL SURFACE DETENT AIL DETENT FOREIGN
COMPARATOR SERVO ENGA MY CAMOUT
ELEV SERVO
LVDT MY DETENT
POSITION DETENT LOGIC
ELEV SURFACE LOW IF LVDT(S) ELEV DETENT TO OTHER CHANNELS
DIFFER BY TO AFDS
RUD SERVO >3` (AIL & RUD) ANNUNCIATIONS
>4` (ELEV) DC RUDDER
RUD SURFACE RUD DETENT ARM RELAY
(FROM RUDDER ARM LOGIC) A
RUDDER ARM
2 IN COMMAND NON-ISO
3 IN COMMAND DISCONNECT
TWO SAME INNER LOOP FAILURE DC
B RUDDER
ENGA RELAY
POWER SUPPLY VALID
TWO SAME AUTOLAND
SENSORS FAILURE (FROM RUDDER ENGA LOGIC) RUDDER ENGAGE
FOREIGN SEVO ENGA
FLIGHT CONTROL COMPUTER (TYP)
1 INHIBIT LOGIC LOW DURING 2 USED DURING DUAL CHANNEL OPERATION ONLY - IF FOREIGN CHANNEL
POWER UP TEST ONLY CAMOUT YOU GET A DETENT TRIP (DETENT LOGIC LOW)

ENGAGE LOGIC DETAILS - HARDWARE MONITORS

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INTERFACE - FCC CROSS-CHANNEL DATA

General

Digital data is transmitted cross-channel between FCCs for monitor, signal

selection, and synchronization. The analog discrete logic for servo engage and
detent trip also go on the cross-channel. They go to dedicated hardware to
provide a redundant path for failed FCC disengagement.

Digital Interfaces

The digital data buses are ARINC 429 high speed buses. Cross-channel bus
parameter labels generally are not identical to the respective source label. This
allows a simplified, combined cross-channel receive function.

Analog Discrete Interfaces

These discretes are either ground or 28v dc.

 SERVO ENGA RELATIVE
DETENT TRIP LEFT 1

SERVO ENGA RELATIVE

RIGHT X CHANNEL
DETENT TRIP RIGHT
X CHANNEL
LEFT
SERVO ENGA
MY DETENT TRIP
LEFT FCC

SAME AS LEFT FCC

1 CROSS-CHANNEL DATA
CENTER FCC
AILERON COMMAND
ELEVATOR COMMAND
RUDDER COMMAND
1
DISCRETE WORDS
GENERAL PURPOSE WORDS

IRS SIGNALS SAME AS LEFT FCC

ADC SIGNALS
ILS SIGNALS
FMC SIGNALS
RADIO HEIGHT
FLAP POSITION RIGHT FCC
STAB POSITION

AFDS GENERAL - INTERFACE - FCC CROSS-CHANNEL DATA

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AFDS POWER DISTRIBUTION Autoland Status Annunciators (ASA) DC Power

General Distribution Each ASA uses two sources of dc power to provide the necessary power
redundancy.
The power distribution for the AFDS is organized around the three channel
configuration of the flight control computers, the dual configuration of the mode
control panel, and the dual requirements of the annunciation and caution and
warning designs.

Each FCC gets power from a separate ac and dc bus. The dc standby bus
provides the second source of annunciation and warning power to each FCC
and to each autoland status annunciator.

AFDS AC Power

The 115v ac primary power to each FCC is changed to 26v ac and used for
sensor excitation and to make several dc levels for the electronic circuits. An
automatic reset circuit breaker protects the 26v ac output. Lighting power (5v
ac) goes to the AFCS mode control panel and each autoland status annunciator
for background lighting.

FCC DC Power

Two external dc sources go to each FCC. The AFCS warning power dc and that
developed from the 115v ac power provide the dual sources necessary for
annunciation and warning.

AFCS MCP DC Power

The AFCS MCP uses two sources for logic power and two sources for master
dim and master test of the indicator lights. Each FCC SERVO dc goes through
the AFCS MCP to allow positive disengagement of the servos from the
AFCS MCP.

The left and right logic power dc inputs supply the left and right side AFCS MCP
electronics respectively. The left and right indicator power inputs provide the
power to the left half and right half of each dot bar display.
 28V DC
STBY BUS 28V DC (BAT)
AUTO FLT
WARN DC 5V AC
28V DC PWR
28V DC MODE CONT
BUS PANEL L AFCS WARN PWR CAPT ASA
LEFT
SERVO ARM PWR
L FCC
SERVO SERVO ENG PWR
115V MAIN L LOGIC POWER
AC BUS PRIMARY PWR
LEFT L FCC PWR
L FCC
115V AC L SERVO POWER
BUS
CENTER C FCC PWR
28V DC SAME AS
BUS L FCC
CENTER C FCC SERVO

C FCC
28V DC C SERVO POWER
BUS MODE CONT
RIGHT PANEL

R FCC SERVO SAME AS

L FCC
115V MAIN
AC BUS
RIGHT R FCC PWR
R FCC
R SERVO POWER

5V AC LIGHT PLATE PWR

SAME AS MASTER DIM IND PWR
5V AC L
CAPT ASA MASTER TEST GND FOR TEST
MASTER DIM IND PWR
MASTER TEST R
GND FOR TEST
F/O ASA
R LOGIC POWER
AFCS MCP

AFDS POWER DISTRIBUTION

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AFDS FCC POWER DISTRIBUTION

FCC AC Power

Each autopilot servo receives 26vac for LVDT excitation from its FCC.
 26V AC SERVO POSN LVDT
LVDT EXCITATION SURF POSN LVDT
L LCCA
L FCC SERVO POSN LVDT
SURF POSN LVDT
L ELEVATOR
A/P SERVO
SERVO POSN LVDT
SURF POSN LVDT
L DIRECTIONAL
A/P SERVO

SAME AS 26V AC SAME AS L

COMPONENT
L FCC
C LCCA

C FCC SAME AS L
COMPONENT
C ELEVATOR
A/P SERVO
SAME AS L
COMPONENT
C DIRECTIONAL
A/P SERVO

SAME AS 26V AC SAME AS L

COMPONENT
L FCC
R LCCA

R FCC SAME AS L
COMPONENT
R ELEVATOR
A/P SERVO
SAME AS L
COMPONENT
R DIRECTIONAL
A/P SERVO
AFDS FCC POWER DISTRIBUTION
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AUTOLAND POWER SWITCHING - FUNCTIONAL bus tie breakers (BTBs) go into the isolation mode. In the isolation mode, both
BTBs can not be closed to supply power to the two main buses
DESCRIPTION from one source.

General Operation

It is necessary for the airplane to have Category III B capability. The categories The standby electrical system is an independent power source to supply the
are established by the regulatory agencies based on these factors: center autopilot during autoland operation. During autoland operation, the
battery charger (which operates in the T-R mode as a constant voltage source)
- Airplane certification is the third power source. The battery provides instantaneous backup if the
- Airport facilities battery charger fails.
- Flight crew proficiency
- Airline maintenance. An autoland bus isolate request starts the center bus transfer. When the K122
ISLN REQUEST relay is energized, a 28v dc signal (Autoland CMD) goes to the
The runway visual range (RVR) and decision height (DH) requirements bus power control unit (BPCU). The BPCU sends an autoland lockout command
established for category III B are typically 150 - 700 ft RVR and 0 ft DH. These to the left and right generator control units to inhibit BTB close command logic.
numbers are based on a fail-operational capability. This is the level of At least one BTB must be open and one GCB closed to energize the autoland
redundancy necessary such that if there is a single failure that occurrs below relays (K526 and K527). For backup protection, the circuit to the BTB close coil
the alert height, the landing may be continued with the remainder of the is open to interrupt any BTB close command when the autoland relay is
automatic system. For a fail-operational capabilty, triple redundancy of these is energized.
necessary:
When either K526 or K527 is energized, if the static inverter is operational, relay
- Power sources K2127 energizes. This lets center isolation relay, K123, energized. If the
- Engaged FCCs battery switch is ON, K2128 and K107 (center bus transfer relay) energize
- Sensors switching the center buses. Additional contacts of relay K2128 bypass K2127
- Servos. after isolation occurs. After center bus transfer is complete, the bus isolated
ground input (from K108) is removed when K107 is energized.
After the FCCs send an autoland bus isolate command to the isolation request
relay (K122), they monitor a bus isolation signal.

Normal Configuration

During non-autoland conditions, the left and right generator circuit breakers
(GCB) are closed, the bus tie breakers (BTB) are open, and the K107 center
bus transfer relay is de-energized. The center bus ac and dc is supplied from
the left main ac and dc buses respectively.

Autoland Configuration

When the three autopilots are armed for approach and there are no faults, the
center buses transfer from the left system to the standby system. This provides
three independent power sources for the three autopilots. At the same time, the
 STBY BUS
A
C BUS AC
POWER

115V AC BUS - C
L

MAIN AC BUS - L
GEN
L BTB L GCB
C BUS AC
INV PWR TRU AUTOLAND
STANDBY
INVERTER
BAT. BUS

DC BUS - L
T-R STBY
CONT POWER SW
1 K106 TO AC
APC MAIN BAT. STBY RLY
APU XFR RLY AUTO BAT.
GEN B
OFF C BUS DC POWER
EPC

28V DC BUS - C
EXTERNAL POWER K526 LEFT
AUTOLAND
RELAY C BUS DC
AUTOLAND
CAT 3 BUS ISO TO K137
A AUTOLAND IRS DC
BAT. COMMAND PWR DISC
TO BUS
CAT 3 BUS ISO POWER C BUS
B CONTROL CONT

HOT BAT BUS

LEFT UNIT K107
CENTER
CAT 3 BUS ISO K123
BAT. XFR BUS XFR
C CENTER
CONT ISOL RLY RELAY
RIGHT K122
BUS
ISOL BATTERY ISOLATED
REQUEST K104 SWITCH
RELAY MAIN
AUTOLAND BAT. L FCC
BUS ISOL ON
RLY

2 28V BUS
L FCC
ISOLATED
K108
AUTOLAND DC TIE C FCC
RELAY R MAIN LOW
BUS ISOL DC TIE
CONTROL UNIT F/O INST BUS
K527 RIGHT VOLTAGE SENS BUS
C FCC
AUTOLAND RLY ISOLATED
MAIN AC BUS - R

AUTOLAND DC BUS - R
R FCC
BUS ISOL T-R C
R BTB R GCB
R FCC R L MAIN LOW
GEN
CAPT INST BUS
1 WHEN ENERGIZED CONTACTS NOT SHOWN VOLTAGE SENS
PREVENT L BTB FROM TRANSFERRING

2 WHEN ENERGIZED CONTACTS NOT SHOWN

PREVENT R BTB FROM TRANSFERRING

AUTOLAND POWER SWITCHING - FUNCTIONAL DESCRIPTION

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AUTOLAND POWER SWITCHING - FUNCTIONAL A generator failure causes the generator circuit breaker (GCB) to trip open. This
causes the corresponding autoland relay to de-energize and the BPCU to
DESCRIPTION CONT. directly sense the GCB position. An open GCB causes the BPCU to remove the
autoland lockout command from the corresponding GCU and allows
Bus Isolation Signal
automatic BTB closure. Transfer to the APU generator may result in a power
interruption from less than 0.05 seconds to 10 seconds, depending on the type
The bus isolation sigmal shows that the buses are isolated. All three channels
of failure.
have an independent source of power. The bus isolation condition is removed in
one of these three ways:

- Radio altitude is above 200 feet and a fault has been detected which results in
a NO LAND 3 condition
- After start of an auto go-around when the airplane is above 100 feet radio
altitude and a positive rate of climb has been established
- All autopilots are disconnected.

Generator Loss Above 200 Feet

Between arming of the three autopilots and 200 feet, the loss of a generator is
sensed by either the captain or first officer instrument bus voltage sensing unit.
Either unit energizes to keep the flight instrument transfer bus powered,
energize the dc tie relay (K108), and signals the FCCs that the buses are not
isolated (re-applies ground to bus isolate signal). The FCCs remove the
autoland bus isolate ground. This hanges the center buses back to left main ac
and dc buses. Also the bus tie breakers close to supply power to both main ac
buses.

Generator Loss Below 200 Feet

If a generator is lost below 200 feet radio altitude, the loss of generator is
sensed by either instrument bus voltage sensing unit. Either unit energizes as
previously discussed. Below 200 feet, the FCCs do not remove the autoland
bus isolate request (ground) and K122 remains energized. The three autopilots
remain isolated. Because of the generator loss, one of the autopilots is also lost,
and the automatic approach continues with the remaining two autopilots.

APU Generator Backup Power

If the APU generator is on during a multichannel approach, it supplies power to

the main ac bus of the failed generator, regardless of altitude.
 STBY BUS
A
C BUS AC
POWER

115V AC BUS - C
L

MAIN AC BUS - L
GEN
L BTB L GCB
C BUS AC
INV PWR TRU AUTOLAND
STANDBY
INVERTER
BAT. BUS

DC BUS - L
T-R STBY
CONT POWER SW
1 K106 TO AC
APC MAIN BAT. STBY RLY
APU XFR RLY AUTO BAT.
GEN B
OFF C BUS DC POWER
EPC

28V DC BUS - C
EXTERNAL POWER K526 LEFT
AUTOLAND
RELAY C BUS DC
AUTOLAND
CAT 3 BUS ISO TO K137
A AUTOLAND IRS DC
BAT. COMMAND PWR DISC
TO BUS
CAT 3 BUS ISO POWER C BUS
B CONTROL CONT

HOT BAT BUS

LEFT UNIT K107
CENTER
CAT 3 BUS ISO K123
BAT. XFR BUS XFR
C CENTER
CONT ISOL RLY RELAY
RIGHT K122
BUS
ISOL BATTERY ISOLATED
REQUEST K104 SWITCH
RELAY MAIN
AUTOLAND BAT. L FCC
BUS ISOL ON
RLY

2 28V BUS
L FCC
ISOLATED
K108
AUTOLAND DC TIE C FCC
RELAY R MAIN LOW
BUS ISOL DC TIE
CONTROL UNIT F/O INST BUS
K527 RIGHT VOLTAGE SENS BUS
C FCC
AUTOLAND RLY ISOLATED
MAIN AC BUS - R

AUTOLAND DC BUS - R
R FCC
BUS ISOL T-R C
R BTB R GCB
R FCC R L MAIN LOW
GEN
CAPT INST BUS
1 WHEN ENERGIZED CONTACTS NOT SHOWN VOLTAGE SENS
PREVENT L BTB FROM TRANSFERRING

2 WHEN ENERGIZED CONTACTS NOT SHOWN

PREVENT R BTB FROM TRANSFERRING

AUTOLAND POWER SWITCHING - FUNCTIONAL DESCRIPTION CONT.

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AFDS GENERAL - POWER ISOLATION LOGIC

General

The purpose of the power isolation logic is to cause the power sources to
separate during a triple-channel approach.

The power isolation logic is set when these conditions are true:

- Triple channel arm - all three channels either engaged or armed

- None of the three channels make NO LAND 3 logic - Approach mode armed
- Airplane above alert height (>200 ft RA). Any of these conditons cause the
power isolation logic to reset:
- Total autopilot disconnect
- Any channel makes NO LAND 3 logic above alert height
- Go-around started.

Note: NO LAND 3 status will cause power isolation logic to be reset but
the reset of fail-operative capability will not restart power isolation
logic unless the approach mode or triple arm is disengaged and
reengaged. Through the power isolation circuits, NO LAND 3 logic
will be made if isolation does not occur within 4 seconds of the
request by the FCC and isolation will be reset and not requested
again.
 MY NO LAND 3 OR
NO AUTOLAND
L NO LAND 3 OR
28V
NO AUTOLAND
A
R NO LAND 3 OR
D
NO AUTOLAND
A
MY SVO ENGA K122 ISOLATION
T
REQUEST RELAY
A
MY SVO ARM
TO R FCC
L SVO ENGA TRIPLE M
OS S Q A
L SVO ARM ARM TO C FCC
R N
R SVO ENGA A
G
R SVO ARM E
M
APPROACH ARM E
N
BELOW ALERT HEIGHT
T
(<200' RA)

A
MY DISCONNECT
L DISCONNECT
GA INITIATE
R DISCONNECT TO NO LAND
4 SEC
3 LOGIC
SOFTWARE BUS ISOLATED
HARDWARE

LEFT FLIGHT CONTROL COMPUTER

AFDS GENERAL - POWER ISOLATION LOGIC

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AUTOLAND SEQUENCE

APP

The following is actions of the aircraft’s Autopilot during different phases of

Autoland flight.

• Approach selected on the MCP busses split autopilots not engaged Arm.
Center bus powered from the hot battery bus.
• GA armed on EICAS as a thrust mode. G/S captured, Flaps not up or G/A
button selected.
• 1500 Ft ASA displays true land status. Remaining Autopilots engage. Flare
and Rollout Arm.
• 500 Ft De-crab
• 300 Ft Stab trim bias
• 200 Ft Rising runway. Alert height no one single failure of a system can
cause an ASA status change.
• 45 Ft Flare
• 25 Ft Throttle retard occurs and Idle is displayed on the FMA throttle
sector.

GA

Aircraft climbs 2000 feet per min. Wings level following LOC till a different pitch/
roll mode is engaged other than GA above 400 Feet.
 * When APPROACH MODE is armed and Autopilot is engaged, remaining operational ones arm.
B-767 Autoland and Go Around Center Autopilot powered from HBB and A/C Inverter.
* At LOCALZER CAPTURE Heading Bug aligns with course and ADI LOC scale expands.
(intercept heading must be with 120 degrees of localizer course o capture. Expect some
oscillation at large intercept angles).
* GO-AROUND arms with either flaps out f up or glide slope capture (no ADI annunciation,
but check EICAS thrust mode).
GLIDE SLOPE CAPTURE occurs at about 1/4 dot above (intercept heading must be within
80 degrees of localizer course).
At MISSED APPROACH ALTITUDE:
ALT CAPTURE, then ALT HOLD
(if set into MCP altitude window)
NOTE: Go-Around Pitch and Roll
modes must be disengaed separately.
Above 400’ RA
500’ RA * Engage any Pitch / Roll Mode
* RUNWAY ALIGNMENT (not annunciated) back to original Autopilot)
1500’ (Radio Altimeter)
* ASA: LND 3 or LAND 2
* Remaining Autopilots Engage * Autothrottle reduces to
old 2000 FPM rate of climb. Below 400’ RA
* FLARE & ROLLOUT arm * To disengage GA
* Rudder active for asymmetrical thrust 330’ RA disengage Autopilots and both
* AUTO NOSE UP TRIM (not annunciated) Flight Directors.
200’ RA
ASA: NO LAND 3 inhibited
* Center buses Auo Transfer inhibited

45’ RA
* FLARE annunciates
* Power goes to Go-Around thrust
25’ RA
* Pitches to hold existing speed, “Bug” speed of ALPHA
* IDLE annunciates speed, whichever is higher.
AUTO GO-AROUND * Roll maintains ground track that exists at the time of engagement.
* Above 5’ activate either Go-Around Switch

* At 5’ RA ROLLOUT annunciates
* After5’ RA + 2 seconds
Auto Go-Aroud inhibited

TOUCHDOWN

* Autopilot holds centerline (uses LOC) * Selecting Reverse disengages Autothrottle (no warnings)
* Autopilot holds nse down yoke * Reversers deploy Auto Speed Brakes (if not already deployed)
* FLARE disengages * Disconnect Autopilot to turnoff runway
* Autobrakes should engage

AUTOLAND SEQUENCE
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THRUST MANAGEMENT SYSTEM - INTRODUCTION

General

The thrust management system controls engine thrust with mode requests from
these:

- Thrust mode select panel (TMSP)

- AFCS mode control panel (MCP)
- Flight management computing system (FMCS).

The thrust management computer (TMC) uses these mode requests to

calculate the autothrottle and thrust limit data.

The TMC also does continuous fault data management and on-ground
maintenance tests.

Description

The TMSP supplies thrust limit mode inputs to the TMC for thrust limit
calculations.

The MCP supplies autothrottle mode inputs to the TMC for throttle position data.

The FMC supplies data to the TMC to control the throttles and thrust limits.

Sensor Inputs

The TMC uses airplane, navigation, and engine sensors to calculate

autothrottle and thrust limit data.

Display Outputs

The TMC shows the autothrottle engage status and modes on the EADIs. It also
shows the thrust limit modes and values on EICAS.
 THROTTLE
CONTROL
THRUST MGMT
COMPUTER

TO RIGHT
ENGINE

FMC ELECTRONIC
EICAS AIRPLANE & ENGINE
SERVO AND EFIS NAVIGATION CONTROLS AND
DRIVE DISPLAYS SENSORS SENSORS

TMSP CONTROLS

THRUST MODE SELECT PNL

MCP CONTROLS
AND DISPLAYS

AFCS MODE CONTROL PNL THRUST MANAGEMENT COMPUTER

THRUST MANAGEMENT SYSTEM - INTRODUCTION

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TMS - COMPONENT LOCATION

General

TMS components are in these locations:

- Flight compartment
- Main equipment center.

Flight Compartment Components

The autothrottle ARM and MODE switches are located on the AFCS mode
control panel (P-55).

The A/T disconnect light is on the P1 panel. The light is amber.

The autothrottle disengage switches are on the outside of each throttle lever.

The go-around switches are on the lower aft edge of each throttle lever.

The servomotor is on the left side of the autothrottle drive assembly.

The microswitch pack assembly is on the bottom of the autothrottle assembly.

A leading edge slat switch is located behind an access panel to the right of the
first officr stabilizer trim indicator.

Main Equipment Center Components

The thrust management computer is on the E1-3 shelf.

 AFCS MODE P3 INSTRUMENT PANEL
CONTROL - THRUST MODE
PANEL (P55) SELECT PANEL

LE SLAT SWITCH
P1 INSTRUMENT PNL (BEHIND PANEL)
- A/T DISCONNECT
LIGHT

SERVO MOTOR
(UNDER FLOOR) E1-3 SHELF
MICRO SWITCH PACK - THRUST MGMT
ASSEMBLY COMPUTER
(UNDER FLOOR)
GO-AROUND
SWITCH (2)

AUTOTHROTTLE
DISENGAGE
SWITCH (2)

TMS - COMPONENT LOCATION

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TMS - THROTTLE COMPONENT LOCATION

Autothrottle Servo Motor

The autothrottle servo motor is at WL 200 under the

throttle control stand. It moves the throttle and
cables through the clutch pack assembly.

A/T Disengage Switch

One autothrottle disengage switch is on each throttle

lever.

Go-Around Switch

Go-around switch input to the TMC is from the same

switches used by the left flight control computer.
Switch Pack Assembly
The switch pack assembly contains two sets of switches.
One set is operated by the left thrust lever. The other
is operated by the right. Thrust reverse switches put
the thrust limit in REV mode and disengage the
autothrottle.

Brake Pack Assembly

The servomotor moves the throttle levers through the

brake pack assembly.
 A/T DISENGAGE
SWITCH

GO-AROUND
SWITCH

FORWARD
ACCESS DOOR

THRUST LEVER
CONTROL RODS
BRAKES
AUTOTHROTTLE
BRAKE PACK
ASSEMBLY

THRUST LEVER
AUTOTHROTTLE ANGLE RESOLVER
SERVOMOTOR
TLA RESOLVER
LINKS (2)
THRUST LEVER
ANGLE RESOLVER

TMS - THROTTLE COMPONENT LOCATION

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TMS - THRUST MODE SELECT PANEL The reference thrust mode on EICAS changes from TO to D-TO when the
assumed temperature is the highest of the three. You can set the assumed
temperature from 0C to 70C.
General Description
The maximum derate is 25%.
The thrust mode select panel weighs 1.5 lb (0.7 kg) and measures 2.26 x 3.27 x
6.9 inches. Power comes from the TMC. There is no internal computer. The
panel uses an internal clock to make the ARINC 429 words.

Thrust Limit Mode Selection

You use the TMSP push-button for the necessary manual thrust limit mode. A
push-button for the currently selected mode has no effect.

The TO/GA push-button selects:

- Takeoff (TO) thrust limit mode if you push the TO/GA push-button when the
airplane is on the ground
- Go-around (GA) thrust limit mode if you push the TO/GA push-button when the
airplane is in the air.

Thrust Derate Selection

The thrust management system decreases the thrust limit with a fixed derate
schedule or an assumed temperature. You select fixed derate schedules with
the TMSP derate push-button 1 or 2.

You select an assumed temperature with the temperature selector (TEMP SEL).
When you turn the temperature selector one detent in one direction or the other,
and the flat-rated temperature shows as the selected temperature on EICAS,
TO stays as the mode annunciator.

The flat-rated temperature is the highest temperature at which the engine can
give maximum thrust. The TMC uses the highest of these three temperatures to
calculate the takeoff thrust limit:

- TAT
- Flat-rated
- Assumed.
 SET DERATE
SELECT SWITCHES

THRUST LIMIT MODE

SELECT SWITCHES (4)

TEMPERATURE
SELECTOR

TMS - THRUST MODE SELECT PANEL

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TMS - OPERATION - EADI DISPLAYS

Autothrottle Modes

These are the operation modes that show:

- N1
- SPD
- FLCH
- GA
- IDLE
- THR HOLD.

These are the limit modes that show:

- FLAP LIM
- ALPHA
- SPD LIM.

TEST shows for the MCDP ground test mode.

 AUTOTHROTTLE MODE (GREEN)
N1
ENGAGED MODE
N1
SPD
FLCH
GA
IDLE
TEST
THR HLD

LIMITS
FLAP LIM
ALPHA
SPD LIM

TMS - OPERATION - EADI DISPLAYS

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TMS - OPERATION - ENGINE EICAS DISPLAY - THRUST LIMIT

FUNCTIONS
General

EICAS provides the displays for the TMC thrust limit functions.

Total Air Temperature and Selected Temperature Displays

Primary data source for total air temperature is the TMC. The ADC is the
secondary data source. Source for selected temperature is the TMC digital data
bus. Assumed temperature is selected on the thrust mode select panel which
goes the TMC.

Reference Thrust Mode

TO, GA, CON, CLB, CRZ, or MAN show in green for the mode. D-TO shows for
assumed temperature derated takeoff mode. 1 or 2 follows the mode for other
derated modes, white if preselected and green if active.

Manual Mode

When the manual thrust set knob is pulled out, the reference thrust mode MAN
is provided by the EICAS computer in green. The reference target cursors and
readouts are green and are controlled by the knob. Manual thrust limit is not
used by the TMC.

Reference/Target Cursor and Reference Limit Display

When the manual thrust set knob is pushed in, the cursor primary data source is
the FMC when VNAV is engaged (the cursor secondary data source is the
TMC) and the reference limit display data source is the TMC. The cursor is
magenta when controlled by the FMC and green when controlled by the TMC.
The reference limit display is always green.

Maximum Limit Marker

The primary data source is the EEC. The TMC is the secondary data source
and is used when the EEC data is invalid or the EEC ALTN light is on.
 TOTAL AIR REFERENCE LIMIT
TEMPERATURE DISPLAY
SELECTED
MAXIMUM TEMPERATURE
LIMIT MARKER _________
+ 12c
TAT
+10c REFERENCE
REFERENCE/ 107.1 TO 107.1
THRUST MODE
TARGET CURSOR 20.0 20.0
10 10

6
2
6
2 THRUST DISPLAY
N
1
THRUST POINTER
385 385
PRIMARY THRUST
EGT
PARAMETER

EICAS DISPLAY

DISPLAY COMPUTER BRT THRUST REF SET

ENGINE STATUS EVENT AUTO BOTH MAX IND
RECORD L R L R RESET

PULL MANUAL THRUST

SET KNOB
EICAS CONTROL PANEL
TMS - OPERATION - ENGINE EICAS DISPLAY - THRUST LIMIT FUNCTIONS
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TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK Digital Signal Outputs

DIAGRAM Digital outputs go to these units:
Thrust Control
- MCP - TMS mode status and vertical speed
- EEC - Engine trim
The autothrottle servo and clutch move the throttle system. The thrust lever
- FMC - TMS mode status, engine bleed status, N1, flap position, and
resolvers electrically transmit the throttle lever angle (TLA) to each engine.
temperature selection
Dual electronic engine controls (EECs) control fuel flow and send engine data
- MCDP - Fault, ground test, and configuration data.
back to the TMC.
Analog Input Signals
EICAS Display
Analog inputs are from these units:
The EICAS display shows N1 data and thrust limit modes.
- EEC discrete GE card
EADI
- A/T disconnect and GA switches
- Air ground relays
The autothrottle mode shows on the EADI.
- Flap slat position module
- Leading edge slat switch.
A/T DISC Light
MCP
The A/T DISC light is an amber caution light controlled by the TMC internal
monitors and self-test functions.
The MCP gives this data:
Sensor Inputs
- A/T ARM
- A/T mode selection
These are the sensor inputs:
- IAS/mach selection
- Elevator speed commands.
- Attitude and rate from the IRS
- Air data from the ADC
TMSP
- Airplane configuration discretes.
The TMSP gives thrust limit mode selection and assumed temperature input.
Digital Signal Inputs
FMC
Digital inputs are from these units:
The FMC gives commands to the A/T in VNAV mode. These are the inputs from
- MCP - Mode selection
the FMC:
- EEC - N1, TLA, and EEC status data
- Thrust limit modes
- FMC - Thrust limits from the MCDU, mode data, and performance data
- IAS and mach targets
- ADC - Air data
- N1
- IRU - Inertial reference data
- GMT
- MCDP - Test data.
- A/T mode requests.
 A/T ARM
IAS/MACH
F/D L NAV
IAS
ON OFF

THR
SEL
V NAV AUTO- R ENGINE
OFF
B THROTTLE EEC DATA EEC TRIM
SPD FL CH

FUNCT C DUAL CHANNEL

ELECTRONIC
MCP TO R ENGINE CONTROL
NI REF ENGINE
FMC
SERVO
THRUST CLUTCH RESOLVER
LIMIT A SYNCHROS
BLEED CONFIG FUNCT
TO
CARDS FUEL
2
GA CLB 1
METERING
TEMP SEL
L-PIMU
CON CRZ POWER LEFT ENGINE
TMSP
INPUT R
DFDAU C
SIGNAL
MANAGEMENT L
EADI
EFIS SYMBOL
B GENERATORS
ADIRU(S) EEC TRIM
C
CONTROL EICAS
COMP
A/G RELAY INTERNAL
MONITOR
AND A/T DISC
L FSPM a
SELF-TEST
LE SLAT SW
28V DC
STBY BUS COLLINS
TAT + 12C
+38C

AFCS WARN A 12
96.1

54.9
D-TO 96.1

26.1
115V AC GRD TEST SEL
UP YES/ADV 8
4
8
115.2
4

L MAIN BUS N1

TMC AC POWER DOWN ON/OFF FLT FAULTS GND TEST NO/SKIP

318 320
28V DC SUPPLY
L MAIN BUS EGT

TMC SERVO
28V DC
L MAIN BUS
TMC
P11 OVERHEAD CB PANEL TMC MCDP EICAS DISPLAY UNIT

TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK DIAGRAM

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TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK

DIAGRAM CONT.
Autothrottle Functions

Thrust control uses N1. N1 protection uses the calculated N1 REF. MCP speed
selection or the FMC controls speed. Rate of climb/descent is controlled in FL
CH. The altitude change gives the vertical speed. Vertical speed is also
controlled in GA. The throttles
come back to the aft stops in the flare mode.

EEC trim control gives trim signals to the selected EEC to decrease small
differences in N1.

Thrust Limit Functions

The thrust limit section uses this data to calculate THRUST LIMIT REF and
MAX N1:

- Engine type
- Mode
- Flight conditions
- Aircraft configuration.

Input Signal Management

The input signal management gives signal source selection and validity monitor
for the digital and analog input signals.

Internal Monitors and Self-Tests

The internal monitors and self-tests give disconnect logic. Recorded faults go to
the MCDP.

BITE

BITE does ground tests requested by the MCDP and sends test results to the
MCDP.
 A/T ARM
IAS/MACH
F/D L NAV
IAS
ON OFF

THR
SEL
V NAV AUTO- R ENGINE
OFF
B THROTTLE EEC DATA EEC TRIM
SPD FL CH

FUNCT C DUAL CHANNEL

ELECTRONIC
MCP TO R ENGINE CONTROL
NI REF ENGINE
FMC
SERVO
THRUST CLUTCH RESOLVER
LIMIT A SYNCHROS
BLEED CONFIG FUNCT
TO
CARDS FUEL
2
GA CLB 1
METERING
TEMP SEL
L-PIMU
CON CRZ POWER LEFT ENGINE
TMSP
INPUT R
DFDAU C
SIGNAL
MANAGEMENT L
EADI
EFIS SYMBOL
B GENERATORS
ADIRU(S) EEC TRIM
C
CONTROL EICAS
COMP
A/G RELAY INTERNAL
MONITOR
AND A/T DISC
L FSPM a
SELF-TEST
LE SLAT SW
28V DC
STBY BUS COLLINS
TAT + 12C
+38C

AFCS WARN A 12
96.1

54.9
D-TO 96.1

26.1
115V AC GRD TEST SEL
UP YES/ADV 8
4
8
115.2
4

L MAIN BUS N1

TMC AC POWER DOWN ON/OFF FLT FAULTS GND TEST NO/SKIP

318 320
28V DC SUPPLY
L MAIN BUS EGT

TMC SERVO
28V DC
L MAIN BUS
TMC
P11 OVERHEAD CB PANEL TMC MCDP EICAS DISPLAY UNIT

TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK DIAGRAM CONT.

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MAINTENANCE FUNCTIONS

General

The following test numbers listed below can be run from the Maintenance
Control Display Panel.

Test 01 is the primary test for the FCC. Test 30 and 40 are required to perform
the Autoland return to service test.

Note: Test 01 is part of test 40.

 MAINTENANCE
CONTROL AND
DISPLAY PANEL
GROUND FLIGHT/
TESTS GROUND FAULTS
LRU GROUND TESTS SYSTEM GROUND TESTS

TEST 04 MCP TEST 30 CURRENT FAULT REPORT

MODE CONTROL
PANEL
TEST 40 AUTOLAND

AUTOLAND TEST 06 ASA

STATUS SUPPORT GROUND TESTS
ANNUNCIATOR
TEST 51 AIR/GND RLY

AILERON TEST 07 SERVO AIL

SERVOS TEST 56 FCC CONFIG/OPT

FLIGHT
CONTROL TEST 59 FCC INSTR
ELEVATOR TEST 08 SERVO ELEV
COMPUTER
SERVOS
TEST 65 STAB TRIM

RUDDER TEST 09 SERVO RUDDER

SERVOS TEST 66 XDCR OUTPUTS

TEST 67 AIL SURF LIM

AUTOPILOT TEST 11 SW A/P DISC
DISCONNECT
SWITCHES TEST 68 ELEV SURF LIM

GO-AROUND TEST 13 SW GA TEST 69 RUD SURF LIM

SWITCHES
TEST 01 FCC

MAINTENANCE FUNCTIONS
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MAINTENANCE MONITOR SYSTEM - REMOTE MCDP

OPERATION
General

There are maintenance operations that can be done by one person with a
remote MCDP control panel in the flight compartment. Annunciation is on the
EICAS display. A hand-held, carry-on panel connects to a plug behind the P6
main power distribution panel. Open the panel to get access. There is a cutout
for the cord.

Description

The remote panel provides the same switch functions as on the front of the
MCDP. Mode annunciations and alphanumeric messages show on the EICAS
display. Select the CONF/MCDP switch on the EICAS maintenance panel
(P61) to show the MCDP display. The airplane must be on the ground to show
the EICAS maintenance pages.
 EICAS MAINT
DISPLAY SELECT

CONF
MCDP

EICAS MAINTENANCE PANEL

MCDP
DISPLAY
AREA
MCDP
UP YES/ADV GND TEST ON/OFF (11 LINES)

DOWN NO/SKIP FLT FAULTS

TYPICAL CARRY-ON MCDP EICAS LOWER EICAS DISPLAY UNIT

REMOTE MCDP CONTROL PANEL COMPUTERS
P6-5
PANEL
PLUG

MAINTENANCE MONITOR SYSTEM - REMOTE MCDP OPERATION

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MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS Each flight control computer gets sensor data from its inertial reference unit and
air data computer on ARINC 429 data buses.
Primary Interface Systems
The thrust management computer gets control signals from the thrust mode
These are the primary interface systems: select panel on an ARINC 429 data bus.

- Flight control computers The thrust management computer supplies analog control signals to the throttle
- Flight management computers servo motor and gets analog servo position signals.
- Thrust management computers
- EICAS computers Ground Test Interface
- Remote MCDP control panel
- Air/Ground relays The primary systems are directly monitored by the MCDP. During the ground
. test, these systems interface with the secondary systems and various sensor/
The left, center, and right flight control computers use ARINC 429 data buses control inputs.
and analog discrete lines. Information includes ground test control and fault
data.

The left and right flight management computers use ARINC 429 data buses to
send fault data. The thrust management computer uses ARINC 429 data
buses and analog discrete lines. Information includes ground test control and
fault data.

The EICAS computers and the MCDP remote control panel let flight
compartment operation of the MCDP.

Secondary Interfaces

These components interface with the primary systems and may be tested from
the MCDP:

- AFCS mode control panel (MCP)

- Servos
- Air Data/Inertial reference units (ADIRUs)
- Thrust mode select panel (TMSP)
- Throttle servo motor.

The mode control panel supplies control signals and gets status data from all six
computers on ARINC 429 data buses.

Each flight control computer supplies analog control signals to its control
surface servos and gets analog servo position signals.
 AFCS MODE CONTROL PANEL

THROTTLE
SERVO
MOTOR GEN

A/T DISC SW (2)

THRUST MODE
AIR DATA/INERTIAL REF UNITS (3) SELECT PNL PACK SWITCHES
ELEVATOR, AILERON AND AUTOPILOT DISCONNECT
RUDDER ACTUATORS (3) ANTI-ICE SWS
SWITCHES (2)
AIR-DRIVEN PUMP SW
STAB TRIM DRIVES (2) GO-AROUND SWITCHES (2)
ISLN VALVE SW
SLAT SWITCHES
AUTOLAND STATUS REV THRUST SW
ANNUNCIATORS (2) FLAP TRANSDUCERS
SLAT SWITCHES
HYDRAULIC VALID OUT
FLAP TRANSDUCER

REMOTE MCDP
AIR/GROUND FLIGHT CTRL FLIGHT MGMT THRUST MGMT EICAS
CONTROL
RELAYS (2) COMPUTERS (3) COMPUTERS (2) COMPUTER COMPUTERS (2)
PANEL PLUG

MAINTENANCE CONTROL AND DISPLAY PANEL

MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS

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MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS These are ARINC 700 sensors:
CONT. - ILS
- ADIRU
Primary and Secondary Systems - VOR
- IRS
Faults in these LRUs can be reported by the MCDP: - DME
- Radar altimeter.
- FCC (3)
- CDU (2) Faults in these LRUs/systems are not reported:
- FMC (2)
- TMC - Automatic direction finding (ADF) system
- AIR/GND relays (2) - Communication radios
- TMSP - Air traffic control (ATC) system
- MCP. - Weather radar
- Caution/warning system
System Interface - Ground proximity warning system (GPWS)
- M/ASI
The AFDS interfaces with these components: - Standby Altimeter
- Yaw damper/YSM
- Flap position - Electronic flight instrument system (EFIS)
- Stab position - Vertical speed indicator (VSI)
- ASA indicators - RDMI.
- A/P servos.

The FMCS interfaces with these components:

- Fuel flow indicator

- Fuel quantity
- Digital clock
- EFIS control panel.

The TMS interfaces with these components:

- Control switches
- Throttle Servo Motors
- PLA transducer
- FSEU
- Flap POS
- EEC.
 AFCS MODE
AFCS MODECONTROL
CONTROLPANEL
PANEL

THROTTLE
THROTTLE
SERVO
SERVO
MOTOR
MOTOR GEN
GEN

A/TA/T DISC
DISC SWSW
(2)(2)
THRUSTMODE
THRUST MODE
AIR
AIR DATA COMPUTERS
DATA/INERTIAL REF (2)
UNITS (3) SELECTPNL
SELECT PNL PACK
PACK SWITCHES
SWITCHES
ELEVATOR,AILERON
ELEVATOR, AILERONAND
AND AUTOPILOT DISCONNECT
AUTOPILOT DISCONNECT
RUDDERACTUATORS
RUDDER ACTUATORS(3)(3) ANTI-ICE SWS
ANTI-ICE SWS
SWITCHES (2)
SWITCHES
AIR-DRIVEN
AIR-DRIVENPUMP
PUMP SWSW
STABTRIM
STAB TRIMDRIVES
DRIVES(2)
(2) GO-AROUND SWITCHES
GO-AROUND SWITCHES (2)
(2)
ISLN VALVE
ISLN SWSW
VALVE
SLAT
SLAT SWITCHES
AUTOLANDSTATUS
AUTOLAND STATUS REVREVTHRUST
THRUSTSWSW
ANNUNCIATORS (2) FLAP TRANSDUCERS
FLAP TRANSDUCERS
ANNUNCIATORS (2) SLAT SWITCHES
HYDRAULIC VALID SLAT SWITCHES
HYDRAULIC VALID OUT
OUT
FLAP
FLAPTRANSDUCER
TRANSDUCER

REMOTE MCDP
REMOTE MCDP
AIR/GROUND
AIR/GROUND FLIGHTCTRL
FLIGHT CTRL FLIGHT
FLIGHTMGMT
MGMT THRUST
THRUSTMGMT
MGMT EICAS
EICAS
CONTROL
CONTROL
RELAYS(2)
RELAYS (2) COMPUTERS (3)
COMPUTERS (3) COMPUTERS
COMPUTERS(2)
(2) COMPUTER
COMPUTER COMPUTERS (2)(2)
COMPUTERS
PANEL PLUG
PANEL PLUG

MAINTENANCE CONTROL AND DISPLAY PANEL

MAINTENANCE CONTROL AND DISPLAY PANEL

MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS CONT.

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MAINTENANCE CONTROL AND DISPLAY PANEL The MCDP ground test mode select switch is a momentary switch/light with a
lamp replaceable from the front panel. It comes on white when engaged and
Purpose commands the CPU to enter the ground test routine.

The purpose of the MCDP is to get and store fault data from the flight GND TEST SEL UP Switch
management system computers and to initiate, select, and process ground test
functions. The ground test select up switch does not have a light. It moves test numbers
up one number each time it is pushed. When you hold the switch in, it slews the
Power Required test numbers up at the rate of five numbers per second.

The MCDP uses 115v ac, 400 Hz, single-phase power. GRD TEST SEL DOWN Switch

Physical Features The ground test select down switch does not have a light. It moves the test
numbers down one number each time it is pushed. When you hold the switch in,
The MCDP is a 6 MCU. it slews the test numbers down at the rate of five numbers per second.

The weight is 15 pounds. YES/ADV Operation Switch

The MCDP uses main equipment center rack air for cooling. The yes/advance switch does not have a light. It is used by the operator to
respond to interrogative messages in the flight fault mode and the ground test
Front Panel Features - Display Window mode.

The MCDP display has two lines of data with 16 characters on each line. Each NO/SKIP Operation Switch
character contains 17 segment gas discharge lamps, including decimal point.
All displays are in the English language. The no/skip switch does not have a light. It is used by the operator to respond to
interrogative messages in the flight fault mode and the ground test mode.
Control Switches - ON/OFF
Operating Instruction Plate
The MCDP ON/OFF switch is a momentary red switch/light with black letters
(FAIL). The lamp is replaceable from the front panel. It starts a power-up routine The instruction panel is on the left side of the front panel. It contains instructions
in the central processing unit (CPU). The power-up routine turns on the switch/ for these functions:
light. The switch/light goes off and the FAIL message is removed after five
seconds if the MCDP passes the self-test. - Apply power and initialize unit
- Select mode of operation
MODE SWITCHES - Respond to interrogative messages
- Ground test operation.
The MCDP flight faults mode select switch is a momentary switch/light with a
lamp replaceable from the front panel. It comes on white when engaged and
commands the CPU to enter the last flight faults routine.
 INFLIGHT
MESSAGE FAULTS MODE
DISPLAY SELECT SWITCH
WINDOW GROUND TEST
Collins MODE SELECT
SWITCH
GROUND TEST
NUMBER SELECT GND TEST SEL
UP YES/ADV YES/ADVANCE
SWITCH (UP SLEW) SWITCH

GROUND TEST POWER MODE

NUMBER SELECT DOWN ON/OFF FLT FAULTS GRD TEST NO/SKIP NO/SKIP
SWITCH (DOWN SLEW) SWITCH
FAIL
r w w

POWER ON/OFF INSTRUCTIONS GROUND TESTS

SELECT SWITCH 1 CYCLE POWER OFF/ON TO SELF TEST.
LRU SUPPORT
VERIFY SEGMENTS LIGHT. VERIFY FAIL 01 FCC 10 SERVO A/T 51 AIR/GRD RLY
LIGHT OUT IN 5 SECONDS. 02 TMC 11 SW A/P DISC 52 TMC RLY/SW
04 MCP 12 SW A/T DISC 56 FCC CONFIG/OPT
2 SELECT MODE (FLT FAULTS OR
05 TMSP 13 SW G/A 57 TMC CONFIG/OPT
GRD TESTS).
06 ASA 14 XDCR COL L 59 FCC INSTR
3 ANSWER QUESTION “?” WITH 07 SERVO AIL 15 XDCR COL R 60 TMC INSTR
YES OR NO SWITCH. 08 SERVO ELEV 16 XDCR WHL 64 SPD BRK/FLAP XDCR
INSTRUCTION PLATE 4 FLT FAULT DATA: LAST FLIGHT 09 SERVO RUD 17 PVD/PVDC 65 STAB TRIM
66 XDCR OUTPUTS
FAULTS/PREVIOUS FLIGHT FAULTS.
67 AIL SERVO LIMIT
5 GROUND TESTS AVAILABLE AS SYSTEM 68 ELEV SERVO LIMIT
SHOWN: (NOTE: FAILURE EXISTS WHEN 30 CURRENT FAULT REPORT 69 RUD SERVO LIMIT
MESSAGE WILL NOT ADVANCE) 40 AUTOLAND

MAINTENANCE CONTROL AND DISPLAY PANEL

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MAINTENANCE CONTROL AND DISPLAY PANEL CONT. Mode Change Routine

Ground Test Listing A mode change routine controls these functions:

The list of ground tests is on the right side of the front panel. It contains test - Auto on
numbers and names for the LRU, SYSTEM, and SUPPORT tests that the - Power up self-test
operator can select. - Flight faults
- ground test routines.
Internal Circuits - Power
Fault Monitor
The power supply provides internal DC power. It contains +5v dc keep alive
power during dormant periods when the airplane is in the air. Fault monitor occurs during the power-up self-test.

I/O Buses Fault Memory

The MCDP gets information from the three flight control computers (FCC), 2 The flight faults are stored in an EPROM nonvolatile memory. The memory can
flight management computers (FMC), and the thrust management computer on be cleared in the shop with ITS test equipment.
digital ARINC 429 receivers. It sends information to the same computers
on ARINC 429 transmitters.

Analog Discrete

Analog discrete output buffer circuits supply four computers with analog ground
test discrete signals.

Software - Executive

The computer software program organization is controlled by an executive

routine that manages these routines:

- Power-up
- Power interruptions
- Mode change
- Periodic functions
- Power down.
 INFLIGHT
MESSAGE FAULTS MODE
DISPLAY SELECT SWITCH
WINDOW GROUND TEST
Collins MODE SELECT
SWITCH
GROUND TEST
NUMBER SELECT GND TEST SEL
UP YES/ADV YES/ADVANCE
SWITCH (UP SLEW) SWITCH

GROUND TEST POWER MODE

NUMBER SELECT DOWN ON/OFF FLT FAULTS GRD TEST NO/SKIP NO/SKIP
SWITCH (DOWN SLEW) SWITCH
FAIL
r w w

POWER ON/OFF INSTRUCTIONS GROUND TESTS

SELECT SWITCH 1 CYCLE POWER OFF/ON TO SELF TEST.
LRU SUPPORT
VERIFY SEGMENTS LIGHT. VERIFY FAIL 01 FCC 10 SERVO A/T 51 AIR/GRD RLY
LIGHT OUT IN 5 SECONDS. 02 TMC 11 SW A/P DISC 52 TMC RLY/SW
04 MCP 12 SW A/T DISC 56 FCC CONFIG/OPT
2 SELECT MODE (FLT FAULTS OR
05 TMSP 13 SW G/A 57 TMC CONFIG/OPT
GRD TESTS).
06 ASA 14 XDCR COL L 59 FCC INSTR
3 ANSWER QUESTION “?” WITH 07 SERVO AIL 15 XDCR COL R 60 TMC INSTR
YES OR NO SWITCH. 08 SERVO ELEV 16 XDCR WHL 64 SPD BRK/FLAP XDCR
INSTRUCTION PLATE 4 FLT FAULT DATA: LAST FLIGHT 09 SERVO RUD 17 PVD/PVDC 65 STAB TRIM
66 XDCR OUTPUTS
FAULTS/PREVIOUS FLIGHT FAULTS.
67 AIL SERVO LIMIT
5 GROUND TESTS AVAILABLE AS SYSTEM 68 ELEV SERVO LIMIT
SHOWN: (NOTE: FAILURE EXISTS WHEN 30 CURRENT FAULT REPORT 69 RUD SERVO LIMIT
MESSAGE WILL NOT ADVANCE) 40 AUTOLAND

MAINTENANCE CONTROL AND DISPLAY PANEL CONT.

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MAINTENANCE MONITOR SYSTEM - COMPONENT

LOCATION
Maintenance Monitor System Component Location

The MCDP is on the E1-2 shelf in the main equipment

center.

The portable remote panel connector is in the P6 panel

behind or below the circuit breaker door panel.

When selected, the MCDP output shows in the MCDP

display area of the lower EICAS display.
 P6 PANEL
- CONNECTOR FOR PORTABLE
REMOTE PANEL

E1-2 SHELF
- MCDP

MAINTENANCE MONITOR SYSTEM - COMPONENT LOCATION

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REMOTE MCDP CONTROL PANEL CONNECTOR LOCATION

Purpose

The remote MCDP control panel connector gives the

connection for a carry-on remote MCDP control panel.

The panel is for ground maintenance only and lets you

operate the MCDP from the flight compartment.

Connector Location

The connector is in the P6 part of the main power

distribution panel.
 CONNECTOR FOR
REMOTE P ANEL (P6)

CONNECTOR FOR REMOTE

MCDP CONTROL PANEL

REMOTE MCDP CONTROL PANEL CONNECTOR LOCATION

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LRU FAULT CONSOLIDATION

Purpose

One of the advanced features available with a system level maintenance

concept
is fault consolidation, where a failure indication may be examined on a system
basis to determine if a sensor or a wiring interface has failed.

LRU Failure

The upper diagram illustrates a failure case of the right inertial reference unit
which is isolated via the indicated logic equation.

Interface Failure

In the lower diagram, a wiring fault is indicated where the input to the FCC R
has been interrupted. By examining the condemnation status on a system
basis, the interface between the right IRU and FCC R is identified as faulty. This
will require deferred maintenance action but prevents the erroneous removal of
either the inertial reference unit or the right flight control computer.

Message Format

These fault consolidation routines are used to determine the bottom line portion
of flight fault messages.
 FAILED FAULT DATA
IRU FCC R
IRU R

FMC L
01 NO LAND 3
IRU R
FMC R

TMC
MCDP
FCC R (FMC L + FMC R + TMC) = IRU R FAILURE
RIGHT IRU FAILURE
FAILED INTERFACE
FAULT DATA
FCC R
IRU R

FMC L
01 NO LAND 3
IRU R/FCC R
FMC R

TMC
MCDP
FCC R IRU R FAILURE = IRU R/FCC R
RIGHT IRU/FCC INTERFACE FAILURE

LRU FAULT CONSOLIDATION

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FAULT MESSAGE FORMATS Ground Test Diagnostics

The diagnostic is displayed in the ground test mode (Test 30 CURRENT FAULT
FLIGHT FAULT MESSAGE FORMAT
REPORT only) by pushing the GRD TEST switch when an LRU fault is being
displayed.
Historical Flight Tag and Fault Storage
Content
The historical flight tag (HFT) is incremented during AUTO POWER ON if, and
only if, the MCDP receives a STORE DATA Bit from one or more FCCs, FMCs
The diagnostic code defines the monitor within a computer that tripped to
and TMC. The STORE DATA Bit is transmitted by one of these computers only
produce the displayed flight deck effect and faulty LRU. The decimal codes and
when the following sequence of events has occurred: IRS ground speed 100
their related monitors are defined in the following tables. An intermittent fault is
KNTS and air/ground = in air, THEN IRS ground speed 40 KNTS and air ground
defined as a fault that sometime in the past, during flight, caused a monitor to
= on ground ,Additionally, an engaged FCC channel will delay setting its
trip, but at some later time during the flight did not. No intermittent bit is present
STORE DATA bit until the channel is disengaged.
for ground test diagnostics.
Number Allocation

The historical flight tag is defined as follows: 00 PRESENT FLT FAULTS - 01

LAST FLT FAULTS - 02-99 PREVIOUS FLT FAULTS

Faulty LRU

The bottom line of the MCDP flight fault message is the faulty LRU or group of
LRU's. The fault information provided on the bottom line of the display should
be sufficient for most unscheduled maintenance activities. A faulty LRU group
message consists of two LRU's with a slash between. The slash mark indicates
wiring may be at fault. The group message is used when the fault can not be
definitely isolated.

DIAGNOSTIC CODES - FORMAT

Description

The MCDP decodes diagnostic fault data received from the FCC, FMC and
TMC computers and provides for a display of a diagnostic message in the
shown format
.
Flight Fault Diagnostics

The diagnostic is displayed in the flight faults mode by pushing the FLT FAULTS
switch when an LRU fault is being displayed.
 FLIGHT FAULT MESSAGE FORMAT

CHANNEL INTERMITTENT OR SSFD

L=1 1 = YES (CASCADING) R INTERMITTENT OR
R=2 XXXXXX 0 = NO (VOTE OUT) SSFD
C=3
EXAMPLE : 121541

SOURCE DECIMAL CODE FCC AIL SRVO SYNC

FCC = 1 ASSIGNMENT (FROM TABLE)
FMC = 2
TMC = 3

DIAGNOSTIC CODE FORMAT

HISTORICAL FLIGHT TAG

00 - PRESENT FLIGHT FAULT
01 - LAST FLIGHT FAULT
02-99 - PREVIOUS FLIGHT FAULT FLIGHT DECK EFFECT
OR DIAGNOSTIC

HFT FDE OR DIAGNOSTIC

FAULTY LRU

FAULTY LRU OR TWO LRU(S) SEPARATED

BY A / WHICH MAY MEAN AN INTERFACE FAULT

FAULT MESSAGE FORMATS

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