B767 ATA 22 Student Book

AUTOPILOT FLIGHT DIRECTOR SYSTEM CH 22

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TRAINING MANUAL FOR TRAINING PURPOSES ONLY

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ATA 22 AUTOPILOT FLIGHT DIRECTOR SYSTEM AUTOPILOT FLIGHT DIRECTOR SYSTEM - INTRODUCTION ............4 TMS - GENERAL DESCRIPTION ...........................................................8 MAINT MONITOR SYSTEM - MCDP AUTO PWR ON MODE..............10 FLT CONT/HYD - GENERAL DESCRIPTION ......................................12 FLT CONT/HYD - FLIGHT CONTROLS - INTRO .................................14 COMPONENT LOCATION - FLIGHT CMPTMNT .................................16 FLT CONT/HYD - COMPONENT - HYD SYS CONTROLS ..................18 FLT CONT/HYD - COMPONENT - FLT CNTRL INDICATORS ............20 FLT CONT/HYD - AILERON CONTROL SYSTEM ...............................22 FLT CONT/HYD - ELEVATOR CONTROL SYSTEM ............................24 FLT CONT/HYD - RUDDER CONTROL SYSTEM ................................26 YSM - YAW DAMPER INTRODUCTION ...............................................28 YSM - YAW DAMPER - GENERAL DESCRIPTION .............................30 YSM - YAW DAMPER - COMPONENT LOCATION .............................32 YSM - GENERAL DESCRIPTION .........................................................34 YSM - GENERAL DESCRIPTION - FUNCTIONS .................................36 YSM - FAULT RECORDING ................................................................38 YSM - BITE ...........................................................................................40 YSM - FAULT ANNUNCIATION ............................................................42 YSM - YAW DAMPER - PRE FLT TEST FROM FLT DECK .................44 YSM - STAB TRIM - INTRODUCTION ..................................................46 YSM - STAB TRIM - COMPONENT LOC - FLT DECK .........................48 YSM - STAB TRIM - CMPNT LOC - MEC AND JKSCRW AREA ..........50 YSM - STAB TRIM - GENERAL DESCRIPTION ..................................52 YSM - STAB TRIM - FUNCTIONAL MODE PRIORITY .........................54 STAB TRIM SYS - MNL ELEC STAB TRIM - DESCRIPTION .............56 STAB TRIM SYS - ALT ELEC STAB TRIM - FCTNL DESC .................58 YSM - STAB TRIM - FCTNL DESC - AUTO STAB TRIM... ...................60 YSM - STAB TRIM- FAULT ANN - STAB TRIM ....................................62 YSM - STAB TRIM - FAULT ANN - UNSCHEDULED TRIM .................70 FLT CONT/HYD - GEN DESC - FLIGHTCONTROLS ...........................72 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

AFDS GENERAL - COMPONENT LOCATION .....................................84 MODE CONTROL PANEL - FUNCTIONS ............................................86 AFDS - ANN AND WARN - GEN DESCRIPTION .................................88 AFDS GEN - AUTOLAND STATUS ANNUNCIATORS .........................90 AFDS - FUNC DESC - ASA - DISPLAY SEQUENCE -1 ........................92 AFDS - FUNC DESC - ASA - DISPLAY SEQUENCE - 2 ......................94 AFDC - FUNC DESC - ASA - DISPLAY SEQUENCE - 3 ......................96 AFDS - FUNC DESC - ASA - DISPLAY SEQUENCE - 4 ......................98 MODE CONTROL PANEL INTERNALS .............................................100 ENGAGE LOGIC - SIMPLIFIED .........................................................102 ENGAGE LOGIC DETAILS - HARDWARE MONITORS .....................104 INTERFACE - FCC CROSS-CHANNEL DATA ...................................106 AFDS POWER DISTRIBUTION ..........................................................108 AFDS FCC POWER DISTRIBUTION .................................................110 AUTOLAND POWER SWITCHING - FUNC DESC ............................112 AFDS GENERAL - POWER ISOLATION LOGIC ................................116 AUTOLAND SEQUENCE ...................................................................118 THRUST MANAGEMENT SYSTEM - INTRODUCTION ....................120 TMS - COMPONENT LOCATION .......................................................122 TMS - THROTTLE COMPONENT LOCATION ...................................124 TMS - THRUST MODE SELECT PANEL ...........................................126 TMS - OP- EADI DISPLAYS ...............................................................128 TMS - OP- ENGINE EICAS DISPLAY - THRUST LIMIT FUNC ..........130 TMS - FUNC DESC - SYSTEM BLOCKDIAGRAM ............................132 TMS - FUNCTIONAL DESC - SYS BLOCK DIAGRAM CONT. ..........134 MAINT FUNCTIONS ...........................................................................136 MAINT MONITOR SYSTEM - REMOTE MCDP OPERATION ...........138 MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS .......140 MAINTENANCE CONTROL AND DISPLAY PANEL ..........................144 MAINTENANCE MONITOR SYSTEM - COMP LOC ..........................148 REMOTE MCDP CONTROL PANEL CONNECTOR LOC ................150 LRU FAULT CONSOLIDATION ..........................................................152 FAULT MESSAGE FORMATS ...........................................................154

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 MANAGEMENT SYSTEM (SINGLE)

MANUAL INPUTS

MANUAL INPUTS

FMCS

THRUST CONTROL (THROTTLES)

AUTOPILOT/FLIGHT DIRECTOR SYSTEM (TRIPLE)

YAW DAMPER SYSTEM (DUAL)

MAINTENANCE MONITOR SYSTEM (MCDP)

YAW CONTROL (RUDDER)

MANUAL INPUTS

ROLL CONTROL SYSTEM (AILERONS, SPOILERS)

MANUAL INPUTS

- FLIGHT DIRECTOR COMMANDS - AFDS ANNUNCIATIONS AND WARNINGS

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AUTO STAB TRIM SYSTEM (DUAL)

PITCH CONTROL SYSTEM (STABILIZER AND ELEVATOR)

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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.

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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 GA

CLB

1

2

TEMP SEL CON

TAT + 12c

THRUST LIMIT FUNCTIONS

CRZ

CRZ

THRUST MODE SELECT PNL

FMCS

UPPER EICAS DISPLAY

SPD A/T ARM IAS/MACH F/D ON

L NAV

IAS OFF SEL THR

V NAV

SPD

FL CH

OFF

AUTOTHROTTLE FUNCTIONS

AFCS MODE CONTROL PANEL

ADI

INTERFACING SYSTEMS GO-AROUND SWITCHES AUTOTHROTTLE DISENGAGE SWITCHES

THROTTLE LEVER ANGLES (TLA)

PIMU

EECS

THRUST MANAGEMENT COMPUTER

SERVO DRIVE TACH FEEDBACK AUTOTHROTTLE SERVO

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MAINTENANCE MONITOR SYSTEM - MCDP AUTO POWER ON MODE Initiation 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.

Faults Not Stored 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 show when you do a manual power-up.

FLIGHT FAULTS L FCC

C FCC

R FCC AUTO ON MODE IN PROGRESS - 3 MINUTES OR - ALL FLIGHT FAULTS STORED

TMC

..... ..

MCDP - AUTOMATIC SHUTDOWN

L FMC

AIR/GROUND RELAYS SENSE TOUCHDOWN R FMC MCDP - AUTO POWER ON - STORE FLIGHT FAULTS IF STORE DATA BITS SET

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FOR TRAINING PURPOSES ONLY

FLT CONT/HYD - GENERAL DESCRIPTION Only one actuator gets power supply from only the left hydraulic system.

System

Continuous Pumps

L or R

EDP

L or R

Demand Pumps

ACMP

Center

ACMP 1

Center

ACMP 2

Center

ADP

Center

RAT

Operation Conditions Basic System Pressure System Pressure During Large Loads Basic System Pressure Basic System Pressure (Auto Load Shed) System Pressure During Large Loads Operates 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.

CAUTION: ACCIDENTAL DEPLOYMENT OF THE RAT MAY OCCUR WHILE DOING GROUND TESTS OF THE PITOT SYSTEM. DEACTIVATE THE SYSTEM BY PULLING THE RAT AUTO CIRCUIT BREAKER ON THE P11 PANEL.

LEFT HYD SYS - FLT CONT

CENTER HYD SYS - FLT CONT

AC MP

EDP

AC MP

AC MP

RIGHT HYD SYS - FLT CONT AC MP

EDP

ADP

L E SLATS RAT

T E FLAPS

STAB TRIM L TAIL SOV

L WING SOV LCCA

STAB TRIM C TAIL SOV

C WING SOV LCCA

SPOILER PCA (1-6-12) AIL PCA (LOB-LIB-ROB)

AIL PCA (LIB-RIB)

1

ELEV FEEL CMPTR

R TAIL SOV

LCCA SPOILER PCA (2-7-11)

SPOILER PCA (3-4-5-8-9-10)

AIL PCA (LOB-RIB-ROB)

1 ELEV A/P SERVO

ELEV A/P SERVO

1 ELEV A/P SERVO

ELEV FEEL CMPTR ELEV PCA (LEFT-RIGHT)

ELEV PCA (LEFT-RIGHT)


DIR A/P SERVO

DIR A/P SERVO RUDDER PCA RIGHT YAW DAMPER SERVO

DIR A/P SERVO RUDDER PCA

LEFT YAW DAMPER SERVO

RUDDER RATIO CHANGER

FLT CONT/HYD - GENERAL DESCRIPTION B767-3S2F Page - 13

R WING SOV

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RUDDER PCA 1

- LOB - LEFT OUTBOARD - LIB - LEFT INBOARD - RIB - RIGHT INBOARD - ROB - RIGHT OUTBOARD

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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.

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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 ELECTRIC TRIM SWITCHES

SPEED BRAKE CONTROL LEVER

FLAP/SLAT CONTROL LEVER STABILIZER TRIM SWITCHES

STABILIZER TRIM SWITCHES

A/P DISENGAGE SWITCH AILERON TRIM POSITION INDICATOR

A/P DISENGAGE SWITCH AILERON TRIM INDICATOR P10 QUADRANT STAND

STABILIZER TRIM POSITION INDICATORS

RUDDER TRIM POSITION INDICATOR CAPT RUDDER PEDALS

AILERON TRIM SWITCHES RUDDER TRIM SWITCH

CAPT CONTROL COLUMN

P8 AFT AISLE STAND PANEL

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F/O RUDDER PEDALS F/O CONTROL COLUMN

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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.

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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 DN

S T A B T R I M

AILERON TRIM

6 4 2 0 2 4 6

APL NOSE UP

0 2 4 6 8 10 O F F

RUD

12 14

AILERON TRIM AIL

POSITION INDICATOR (2) (CONTROL ( WHEEL)

STABILIZER TRIM POSITION INDICATORS

ELEV

AIL

EICAS DISPLAY

(2) (P10)

RUDDER TRIM POSITION INDICATOR

15

10

RUDDER TRIM 5 0 5

10 15

NOSE LEFT UNITS NOSE RIGHT

AILERON TRIM SWITCHES

AILERON

LEFT WING DOWN

RIGHT NOSE WING LEFT DOWN

NOSE RIGHT

COMPUTER

DISPLAY ENGINE STATUS

EVENT RECORD

L

AUTO

R

RUDDER TRIM SWITCH AILERON/RUDDER TRIM CONTROL PANEL (P8)

EICAS CONTROL PANEL (P9)

FLT CONT/HYD - COMPONENT - FLIGHT CONTROL INDICATORS B767-3S2F Page - 21

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BRT BRT BAL

THRUST REF SET L

BOTH

R MAX IND RESET

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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.

FIRST OFFICER CONTROL WHEEL

CAPTAIN CONTROL WHEEL

TO RIGHT WING

R LCCA OUTPUT QUADRANT CENTER LCCA

L LCCA R TORQUE TUBE

L LCCA OUTPUT QUADRANT FEEL CENTER AND TRIM ASSY

L TORQUE TUBE R LCCA TO LEFT WING

AILERON PCA (8)

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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.

CAPTAIN CONTROL COLUMN

CONTROL COLUMN OVERRIDE MECHANISM

FIRST OFFICER CONTROL COLUMN

TENSION REGULATOR QUADRANT (2)

ELEVATOR FEEL COMPUTER

OVERRIDE MECHANISM AUTOPILOT PITCH CONTROL SERVO (3)

FEEL AND CENTERING UNIT HORIZONTAL STABILIZER

AFT QUADRANT INTERCONNECT ROD SLAVE CABLE INTERCONNECT

STICK SHAKER (2) SLAVE CABLE QUADRANT (2)

CENTER LINE OF STABILIZER REAR SPAR HINGES

LEFT AFT QUADRANT OUTPUT ARM CONTROL ROD

LOST MOTION AND OVERRIDE DEVICE (2)

POWER CONTROL ACTUATORS (PCA(S) (3 PLACES ON EACH OUTBOARD ELEVATOR) FWD LEFT OUTBOARD ELEVATOR (RIGHT SIDE SIMILAR) OUTBD

AFT QUADRANTS OVERRIDE MECHANISM

LEFT AFT QUADRANT TORQUE TUBE LEFT INBOARD ELEVATOR (RIGHT SIDE SIMILAR)

FLT CONT/HYD - ELEVATOR CONTROL SYSTEM B767-3S2F Page - 25

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POSITION TRANSMITTER (2) RIGHT AFT QUADRANT OUTPUT ARM CONTROL ROD

RIGHT AFT QUADRANT TORQUE TUBE

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FLT CONT/HYD - RUDDER CONTROL SYSTEM Rudder System Description

Trim inputs from the flight compartment control switch move the aft quadrant 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 main equipment center. These components are in the vertical stabilizer:

Power Control Actuators (PCA)

- Autopilot rollout guidance servos - Aft quadrant with the trim feel and centering unit - Ratio changer mechanism - Yaw damper actuators - 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

Three PCAs move the rudder each with a different hydraulic system. The left hydraulic system pressure to the middle PCA goes through the ratio changer actuator. If the ratio changer function has a failure, the middle PCA depressurizes. Each PCA has an override in the input linkage to its control valve.

PCA (3) (C) YAW DAMPER MODULES L AND R

YAW DAMPER SUMMING MECHANISM

(L)

TEMPERATURE COMPENSATION LINKAGE YAW DAMPER SERVOS

RUDDER RATIO CHANGER MODULE

(R) RUD

AIL

ELEV AIL

POSITION TRANSMITTER EICAS DISPLAY STATUS PAGE

RATIO CHANGER ACTUATOR TRIM ACTUATOR

RATIO CHANGER MECHANISM

A

RUDDER PEDALS FWD

A AFT QUADRANT ASSEMBLY FEEL, CENTERING, AND TRIM MECHANISM FLIGHT CONTROL COMPUTERS L,C,R

DIRECTIONAL AUTOPILOT SERVOS (3)

FLT CONT/HYD - RUDDER CONTROL SYSTEM B767-3S2F Page - 27

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RUDDER TRIM SWITCH AND TRIM POSITION INDICATOR

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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 (DUE TO YAW)

LEFT WING INCREASES LIFT & DRAG SIDESLIPS LEFT

WIND

CL

RELATIVE

MOVES LEFT, ROLLS RIGHT

WEATHERVANE MOMENT

YAW ANGLE

TURN WITHOUT RUDDER APPLICATION

BANK STARTS TURN LEFT

RIGHT WING INCREASES LIFT, DRAG ROLLS LEFT, YAWS RIGHT WEATHER COCKS PAST STABLE HDG A/C MOVES RIGHT OF TRACK TAIL MOVES RIGHT OF AXIS SIDESLIPPING OCCURS GUST SIDE FORCE

SIDESLIP

YSM - YAW DAMPER INTRODUCTION B767-3S2F Page - 29

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DUTCH ROLL

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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

R

YAW DAMPER TEST SWITCH

EICAS

FMC YAW DAMPER L R

FSPM

HYD SYSTEM PRESS SWS

YAW DAMPER CONTROL PANEL

AIR

YAW DAMPER/STABILIZER TRIM MODULE (YSM)

YAW DAMPER SERVOS GND AIR/GND SYS

AD IRU

ACCELEROMETERS

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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 - YAW DAMPER CONTROL PANEL

P11 OVERHEAD CIRCUIT BREAKER PANEL - LEFT YDM CB - RIGHT YDM CB - LEFT AND RIGHT FLT CONT ELEC PWR SUPPLY C/B'S MECHANICAL LINKAGE (REF) YAW DAMPER SERVOS

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 B767-3S2F Page - 33

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RUDDER POWER CONTROL ACTUATORS (REF)

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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/ stabilizer trim modules (YSMs). The YSMs use different inputs to give outputs to actuators and modules. 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.

BITE Functions The YSM built-in test equipment (BITE) software give fault messages and details to identify yaw damper and stabilizer trim system component faults.

MANUAL TRIM TEST

ALTERNATE TRIM EICAS

STCM

YAW DMPR L

CUTOUT SWITCHES

R

YAW DAMPER TEST SWITCH

LIMIT SWITCHES

FCC

P5 PANEL FMC YAW DAMPER/STABILIZER TRIM MODULE (YSM)

FSPM

HYD SYSTEM PRESS SWS

YAW DAMPER L R

ADIRU (3) AIR

YAW DAMPER CONTROL PNL

GND AIR/GND SYS

ACCELEROMETERS

YSM - GENERAL DESCRIPTION B767-3S2F Page - 35

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YAW DAMPER SERVOS

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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 STRUCTURAL MODA L SUPPRESSION STRUCTURAL NOTCH FILTER YAW RATE DAMPING AT HIGH ANGLE OF ATTACK SYSTEM D ISENGAGEMENT FAILURE INDICATIONS MAINTENANCE INDICATIONS

CALCULATION OF VC AND MC LIMITS CONTROL WHEEL INHIBIT TO SPOILER CONTROL MODULE REDUNDANCY MANAGEMENT FAILURE INDICATION OUTBOARD AILERON LOCKOUT FUNCTION

YAW DAMPER FUNCTION CALCULATION OF ELEVATOR FEEL TAKEOFF PRESSURE LIMIT FAILURE INDICATION ELEVATOR FEEL LIMIT FUNCTION

RUDDER COMMAND GAIN CONTROL LAW REDUNDANCY MANAGEMENT FAILURE INDICATION MAINTENANCE INDICATIONS RUDDER RATIOCHANGER FUNCTION

NON-VOLATILE STORAGE AND RETRIEVALOF FAILURE AND MAINTENANCE RELATED DATA SELF-TEST MAINTENANCE TEST BITE FUNCTION

AUTOTRIM INTERFACE MANUAL ELECTRIC TRIM INTERFACE MACH TRIM CONTROL LAW STABILIZER TRIM MODE SELECTION MANUAL TRIMRATE MONITORING 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

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YSM - FAULT RECORDING

Flight Leg

General

A new flight leg starts on the transition to takeoff. At takeoff, the faults history 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 memory to store 256 faults and 64 flight legs.

Fault Record

Fault Storage

A fault record is stored for each flight leg in which a fault is recorded. The fault 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 together, the YSMs are the only place to find fault details and identify which YSM detected the fault. Faults can be stored from any of these: - Power-up test - Ground test - Flight phase faults - Internal YSM faults. Power-up test is done at initial power-up of the YSM. The test lasts no more 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 detected, the INOP light and EICAS message stay. 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.

- Fault message - Message number - Flight deck effect (FDE) - Most likely line replaceable units (LRUs) at fault (shown as the reference designator for the LRU) - More possible faulty LRUs or details that cause the fault (if any) - Additional details or wiring diagram (if necessary) - Fault latched or not (not shown in fault history). The fault record(s) for fault history may contain the above details and may include this additional data: - Fault type (hard, intermittent, or repeated) - Flight phase of first occurrence.

FAULTS HISTORY FLIGHT LEG = X

FLIGHT LEG = 0

DESCENT CRUISE

HOLDING EXISTING FAULTS FAULT HISTORY GROUND TEST

CLIMB

APPROACH

POWER-UP TEST GROUND TEST

FAULT MESSAGE

TAKEOFF

MSG NO: XX-YYY

GROUND SPEED > 60 KNOTS

SER

EXISTING FAULTS?

-CAUTION-

LRU: TEXT/* FAULTS HISTORY?

MENU

ON/OFF

YES

NO

GROUND TEST? YAW DAMPER/STABILIZER TRIM MODULE (YSM) P/N 285T1122SERIAL NUMBER MOD LEVEL

D H

C G

B F

A E

OTHER FUNCTIONS?

PERFORM GROUND TEST

(MORE POSSIBLE FAULT LRUS OR DETAILS) FAULT HISTORY ONLY (ADDITIONAL DETAILS, WIRING DIAGRAMS)

YSM - FAULT RECORDING Page - 39

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TYPE:

EXISTING FAULTS GROUND TEST ONLY LATCHED: (YES OR NO)

B767-3S2F

ON GROUND

FDE (TEXT)

REMOVE THE FRONT PANEL SLOWLY TO PREVENT DAMAGE TO INTERNAL CABLE

MFR

LANDING


PHASE:

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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(?) Push to move down in list Push to move up in list Push MENU to return to previous menu

CAUTION OBSERVE PRECAUTIONS FOR HANDLING

ELECTROSTATIC SENSITIVE DEVICES

INSTRUCTIONS:

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 OTHER FUNCTIONS - Shows other functions

BITE MAIN MENU:

LRU RESET:

LRU RESET: Push YES and NO simultaneously for 1 second

EXISTING FAULTS?

YAW DAMPER/STABILIZER TRIM MODULE P/N 285T1122S/N MOD LEVEL

A

B

C

D

E

YAW DAMPER/STABILIZER TRIM MODULE

YSM - BITE B767-3S2F Page - 41

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MENU

ON OFF

YES

NO

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YSM - FAULT ANNUNCIATION General Flight deck effects (FDEs) for the YSM show on these: - Upper and lower EICAS displays - P5 overhead panel - Yaw damper control panel. Upper EICAS Display The message, UNSCHD STAB TRIM (B) shows for any of these: - Stabilizer movement is detected without a valid stabilizer trim command from the YSM - Stabilizer movement opposite to a valid stabilizer trim command from the YSM - Alternate electric trim is used when an autopilot is engaged. The message, STAB TRIM (C) shows if both YSMs command manual trim and only one brake releases. The message, L/R YAW DAMPER (C) shows for any of these: - Channel pin fault - Airplane pin fault - Invalid YSM part number - Yaw damper actuator fault - Yaw damper intermittent - YSM fault (internal fault) - L/R yaw damper disengaged. 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

- Accelerometer fault - Yaw damper acceleration fault (excitation) - FMC fault - Reference 26v ac fault. The message, STAB TRIM (M) shows for any of these: - Channel pin fault - 15 volt power fault - Stab trim control module (STCM) fault - Left and center flight control computer (FCC) (left YSM) - Right and center FCC (Right YSM) - Stabilizer position - Flap 1, 2, and 3 fault - Manual trim switch fault - left hydraulic pressure fault (Left YSM) - Center hydraulic pressure fault (Right YSM) - Stab brake applied - Stab 28v dc fault - YSM fault (internal fault). The message, YSM (S) shows for a YSM internal fault. P5 Overhead Panel The lights on the P5 panel come on for the same reasons as the EICAS messages. Yaw Damper Control Panel 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.

YAW DAMPER L R ON

ON

w

INOP

INOP

a

w

TAT-12C

a

CRZ


WARNING

1.8

UNSCHD STAB TRIM STAB TRIM L YAW DAMPER R YAW DAMPER

1.50

1.8

1.4

1.50

1.00

1.00 1.0

1.4

00.0

1.0

00.0

r

CAUTION

N 1

a

21

MASTER CAUTION/ WARNING LIGHT (2)

21

GET

YSM

YAW DAMPER STAB TRIM

STAB TRIM UNSCHED STAB TRIM

STATUS DISPLAY

a

a

P5 PANEL

ECS/MSG DISPLAY

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YSM - YAW DAMPER - PRE FLIGHT TEST FROM THE FLIGHT DECK General The preflight test uses only the built-in test functions of the yaw damper function. To start the preflight test, push the YAW DMPR test switch on the P61 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 damper servos and the switches on the yaw damper control panel operate correctly. Hydraulics The center hydraulic system is necessary to do a test of the left yaw damper. The left hydraulic system is necessary to do a test of the right yaw damper. Interfacing Systems These systems are necessary during the yaw damper test: - Rudder and rudder trim control system - IRU aligned and in NAV mode - Air/ground relays - Master dim and test system - EICAS. Yaw Damper Systems 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 damper engage switch on the yaw damper control panel and make sure the ON light comes on, and the INOP light goes off. Operation 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 channel in test is on during test of each channel. The yaw damper module starts a test in all three IRUs.

The IRUs transmit these normal IRU test values over the ARINC 429 bus: - Pitch angle - Roll angle - Yaw rate - Roll rate - Lateral acceleration - Ground speed. Because the YSM is in the preflight test mode, it uses these IRU test values and monitors them for correct performance. If an IRU input is not in limits, an L/C/R IRU fault message shows. 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 yaw damper INOP annunciator goes off approximately fifteen seconds after test start to show satisfactory completion of the test. The mode and status annunciators on the inertial reference mode panel come on during the IRU test mode. IRS test values can also show on flight compartment indications. See the air data inertial reference unit section for more information about the IRS test (ATA 34-21). 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 failures, the INOP indication stays on, and the auto disengage relay deenergizes. BCM Preflight Test Indications 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 was a preflight test started from the flight compartment.

WARNING:

KEEP PERSONNEL AND EQUIPMENT AWAY FROM ALL CONTROL SURFACES WHEN HYDRAULIC POWER IS SUPPLIED. INJURY TO A PERSON OR DAMAGE TO EQUIPMENT CAN OCCUR WHEN HYDRAULIC POWER IS SUPPLIED.

ENGAGE YD

PRE FLT YD TEST?

MENU

ON OFF

YES

NO

TURN ON HYD

WARNING! WARNING! L

RUDDER CLEARED?

YD SYS IN TEST CHANNEL TESTED INOP LIGHT ON FOR TIME OF PREFLIGHT TEST YAW DAMPER L R ON

w

INOP a

ON

PASS/FDE DUE TO:

RUD

w

INOP a

AIL ELEV AIL


EICAS STATUS PAGE

YSM - YAW DAMPER - PRE FLIGHT TEST FROM THE FLIGHT DECK B767-3S2F Page - 45

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C

CLEAR RUDDER

TEST PASS

RUDDER INDICATOR

3O

3O

FMC FAULT

R

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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)

CONTROL SWITCH

APL NOSE DN

ARM SWITCH

APL NOSE UP

ALTERNATE ELECTRIC STAB TRIM SWITCHES TO STCM

LIMIT SWITCH AND POSITION TRANSMITTER MODULES STABILIZER TRIM BALLSCREW ACTUATOR ASSEMBLY

STABILIZER POSITION INDICATOR (2)

MANUAL ELECTRIC TRIM SWITCHES

ADC FAIL

COLLINS EXTERNAL SENSOR FAULT

TEST ONLY FUNCTIONAL TEST

MACH DATA TRIM COMMANDS TO STCM THROUGH LIMIT AND CUT-OUT SWITCHES

ADC (2) FLIGHT CONTROL COMPUTER (3)

CAPT CONTROL WHEEL (OUTBD)

AUTOTRIM DATA MANUAL ELECTRIC TRIM CMD

F/O CONTROL WHEEL (OUTBD)

YAW DAMPER/STABILIZER TRIM MODULE (2)

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YSM - STABILIZER TRIM - COMPONENT LOCATION - FLIGHT DECK Control Column Cutout Switches The column cutout switches give manual override of an electric trim command when the control columns move in a direction opposite to the electric trim command. This corrects a runaway trim. P5 Overhead Panel The STAB TRIM light shows if the two YSMs give a manual trim command, and only one brake releases. The UNSCHD STAB TRIM light shows for these conditions: - Stabilizer movement without a valid stabilizer trim command from the YSM - Stabilizer movement opposite to a valid stabilizer trim command from the YSM - 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) .

Stabilizer Trim Control Switches 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 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 moved down gives an airplane nose up trim command, and one or the other set moved up gives an airplane nose down trim command. Alternate Electric Stab Trim Switches Two switches (arm and control) on the P10 quadrant stand give electrical input to the STCMs. Training Information Point Access to the control column cutout switches is through the forward equipment center. The other items are in the flight compartment.

MANUAL ELECTRIC CONTROL SWITCHES

CONTROL COL CUTOUT SW (UNDER FLOOR FIRST OFFICER CONTROL COL)

CONTROL COL CUTOUT SW (UNDER FLOOR CAPTAIN CONTROL COL)

ENTRY DOORS a

EMER DOORS a

CARGO DOORS a

ACCESS DOORS a

CAPT PITOT a L AUX PITOT a

FO PITOT

L AOA

R AOA

STAB TRIM

a

a

R AUX PITOT

a

SPOILERS

a

UNSCHED STAB TRIM a

RUDDER RATIO a

BODY VANE a w

AUTOPILOT CAUTION LIGHT

PULL UP

r

r

CABIN ALT r ALT ALERT a A/T DISC a

AUTO PILOT a FMC

a

r

A/P DISC r OVSPD

w

L ENG PROBE a

R ENG PROBE a

w

w

w

GND

APL NOSE UP

QUADRANT STAND (P10)

a

PROX G/S INHB a

CUT OUT

P1-3 CAPT INST PANEL L

C

STAB TRIM CUTOUT SW

YSM - STABILIZER TRIM - COMPONENT LOCATION - FLIGHT DECK Page - 49

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0 2 4 6


10 12 14

STABILIZER POSITION INDICATOR (2 (2)

STAB TRIM

B767-3S2F

ANTISKID

8

T R I M NORM

w

AIL LOCK a

S T A B

r

a

AUTO SPDBRK a

ANNUNCIATOR PANEL (P5)

ALTERNATE STAB TRIM SWITCHES

CONFIG

a

R TAT

a

APL NOSE DN FIRE

L TAT

a

TAPE

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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 - RIGHT YSM

E1-1 SHELF - LEFT YSM

TRIM LIMIT SWITCH AND POSITION TRANSMITTER MODULES MAIN EQUIPMENT CENTER

STABILIZER JACKSCREW ACCESS AREA

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YSM - STABILIZER TRIM - GENERAL DESCRIPTION General The yaw damper/stabilizer trim modules (YSMs) give these trim commands to the stabilizer trim control modules (STCMs): - Manual trim - Autopilot trim - Mach/speed trim. Manual Electric Trim System The stabilizer trim control switches on the outboard horn of each control wheel give trim up or down arm and control signals to the left and right YSM. The 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 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 assembly forward of the stabilizer. Left and center hydraulic systems supply pressure to the STCMs through STCM cutoff valves. Two hydraulic cutout switches on the top left side of the control stand supply electrical power to the STCM cutoff valve motors. This isolates hydraulic power from the STCMs. Hydraulic inputs from the elevator feel computers control the rate of 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 the flight control computers (FCCs). The position and limit switch modules also control the flight compartment stabilizer position indicators and limit stabilizer travel by cam-operated microswitches that open the electrical path from the YSMs to the STCMs. Column-operated cutout switches below the floor on each outboard side of the control column torque tube also open the electrical path 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 electrical trim limits related to trailing edge flap retract signal from the FSPM.

Autopilot Trim System The three FCCs in the main equipment center supply autopilot input signals to 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 assembly, and limit switch and position transmitter modules is the same as the manual electric system. Mach/Speed Trim System One or the other YSM uses signals from the air data computers (ADC) and 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 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 STCMs, the stabilizer trim ballscrew actuator assembly, and the limit switch and position transmitter modules is the same as the manual electric system. Alternate Electric Stab Trim System The two alternate electric stab trim switches give up or down arm and control trim signals to the dual coil solenoid valves in the STCM. These signals energize one coil in the dual coil solenoid valves (the second coil is controlled by the trim signals from the YSMs). The valves operate as before and supply hydraulic pressure to release the hydraulic brakes and move the hydraulic motors to do a trim of the stabilizer. Stabilizer Trim Fault Indication System The YSMs and FCCs control the logic for the flight compartment amber annunciation on the P5 overhead panel and the indication on EICAS. The YSM also has a BITE control module (BCM) to make a record of the status of system LRUs. Stabilizer Trim Ground Test 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 B767-3S2F Page - 53

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YSM - STABILIZER TRIM - FUNCTIONAL MODE PRIORITY 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.

Manual electric trim inhibits during multi-channel operation. The YSM does not let the manual electric trim inputs disengage the autopilot.

AUTOPILOT DISENGAGED

SINGLE CHANNEL ENGAGED

1 ALTERNATE ELECTRIC OR MANUAL ELECTRIC STAB TRIM

AUTOPILOT

1 ALTERNATE ELECTRIC OR MANUAL ELECTRIC STAB TRIM

2 MACH

YSM - STABILIZER TRIM - FUNCTIONAL MODE PRIORITY Page - 55

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1 ALTERNATE ELECTRIC STAB TRIM 2 AUTOSTAB TRIM

2 AUTOSTAB TRIM

(OPERATION OF MANUAL ELECTRIC TRIM DISENGAGES AUTOPILOT)

B767-3S2F

MULTICHANNEL AUTOPILOT ENGAGED

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(MANUAL ELECT RIC TRIM INHIBITED IN MULTICHANNEL ENGAGE)

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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 VALVE NOSE UP CONT SOLENOIDS

STAB TRIM 28V DC STBY BUS L STAB TRIM SHUTOFF

CUTOUT NORM

28V DC STBY BUS R STAB TRIM SHUTOFF

CUTOUT NORM

CONTROL ARM

P11 CB PANEL

FWD L COL FWD

A B

NOSE UP ARM SOLENOIDS OPEN S/O VALVE NOSE DN CONT SOLENOIDS NOSE DN ARM SOLENOIDS

NOSE UP

FLAPS UP

P10 NOSE DN

FWD R COL FWD L TRIM LIMIT SELECT

ARM

SECONDARY BRAKE RELEASE

AFT R COL AFT

NOSE UP

L STCM

ARM 28V DC STBY BUS L STAB TRIM CONT 28V AC L BUS L CONT STAB TRIM P11 CB PANEL

CONTROL NOSE DN CAPT CONT WHEEL TRIM SWITCH

SAME AS CAPT

ARM

F/O CONT WHEEL TRIM SWITCH

CONTROL

AFT L COL AFT L YSM (E1-1)

SAME AS LEFT

AFT L COL AFT COLUMN CUTOUT SWITCHES

L STAB POS IND

HYD BRAKE

LIMIT SWITCHES POSITION XDUCER L POSN XMTR MODULE

HYD BRAKE

P10

Page - 57

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HYD MOTOR

HYD MOTOR DN

(RIGHT CHANNEL SAME AS LEFT)

R YSM (E2-1)

A

SAME AS LEFT

B

R STCM


UP

STABILIZER BALLSCREW ACTUATOR ASSEMBLY

STABILIZER TRIM SYSTEM - MANUAL ELECTRIC STAB TRIM - FUNCTIONAL DESCRIPTION B767-3S2F

DN

UP

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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 SWITCH CONTROL SWITCH

FLAP

ALT STAB SW FAULT L YSM (E1-1)

APL NOSE DN

APL NOSE UP ALTN STAB TRIM

ENG VALVE a SPAR VALVE a L

REV ISLN a

FUEL CONTROL

ENG VALVE a SPAR VALVE a R

RUN

ALT STAB SW FAULT

CUT OFF

NORM

CUT OUT L

C

NOSE DN CONTROL SOLENOIDS NOSE UP CONTROL SOLENOIDS NOSE UP ARM SOLENOIDS NOSE DN ARM SOLENOIDS SECONDARY BRAKE RELEASE L STCM

STAB TRIM

R YSM (E2-1)

28V DC BAT BUS

IND

UP DN

ARM

UP DN

CONTROL

UP DN

ALT STAB TRIM P11 CB PANEL

HYD HYD BRAKE MOTOR STABILIZER BALLSCREW ACTUATOR ASSEMBLY HYD HYD BRAKE MOTOR DN UP

NOSE UP CONTROL SOLENOIDS NOSE UP ARM SOLENOIDS NOSE DN ARM SOLENOIDS NOSE DN CONTROL SOLENOIDS R STCM

STABILIZER TRIM SYSTEM - ALTERNATE ELECTRIC STAB TRIM - FUNCTIONAL DESCRIPTION Page - 59

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UP

SECONDARY BRAKE RELEASE

ALTN ELEC STAB TRIM SWITCHS

B767-3S2F

DN


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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 AND AUTOTRIM LOGIC

AUTOTRIM VALID TO MCP

1 AUTOTRIM ARM LEFT FCC (E1-1)

TO MCP

SAME AS LEFT FCC CENTER FCC (E1-3)

TO MCP

M A N A G E M E N T

COMMANDS TO STCM

3.5 SEC

AUTOTRIM COMMANDS

AUTOTRIM SELECTION LOGIC

SAME AS LEFT FCC

LEFT YSM (E1-2)

FRAT TRIM TRANSFER LOGIC

TRIM TRANSFER RELAY

SAME AS LEFT SRM

RIGHT YSM (E2-2)

RIGHT FCC (E1-4)

1

SRM FAULT VALID MANUAL COMMAND

D A T A

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 B767-3S2F Page - 61

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YSM - STABILIZER TRIM- FAULT ANNUNCIATION - STAB TRIM General

A YSM fault message that causes a STAB TRIM level C advisory EICAS message and a STAB TRIM annunciator on the P5 overhead panel can show on the EXISTING FAULTS? and FAULT HISTORY? menus. This message is 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 necessary. The YSM fault messages that causes STAB TRIM level M maintenance EICAS message can show on the EXISTING FAULTS? and FAULT HISTORY? menus. 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

- YSM VAL DISCRETE. Stab Brake Switch Fault This fault is caused if the two YSMs give a manual trim command and only one brake releases. Alternate Stab Switch Fault This fault is caused by one of these: - An alternate electric stab trim micro-switch failure - A Mach loop error - stabilizer movement slower than 0.067 degrees/second for 10 seconds of a valid command or 0.3 degrees stabilizer position difference between stabilizer position and commanded position for 10 seconds.

- STCM FAULT - MAN TRIM SW FAULT - L/C HYDR DISCRETE - STAB BRK SW FAULT - STAB 28V FAULT

S Q R

EXISTING FAULTS?

RESET LATCH?

STAB TRIM

a

PANEL P5 OVERHEAD

CSEU +/- 5V DC STAB TRIM 28V DC MENU

YES

ON OFF

NO

- 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 - YSM VAL DISCRETE

TAT

WARNING

STAB TRIM

r

CAUTION a

L/R YSM

EICAS

YSM - STABILIZER TRIM- FAULT ANNUNCIATION - STAB TRIM B767-3S2F Page - 63

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YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB TRIM CONT. STCM Fault The STCM fault is caused by one of these: - Stabilizer movement is slower than 0.067 degrees/second for 10 seconds of a valid command - 0.3 degrees difference between stabilizer position and commanded position for 10 seconds. The fault has a reset 5 seconds after the error goes away or by a stab trim reset. 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.

Channel Pin Fault 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 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 code discretes come back into agreement, the channel code fault has an auto reset. 15v Power Fault 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 reset when the power input comes back in valid range.

- STCM FAULT - MAN TRIM SW FAULT - STCM FAULT - L/C HYDR DISCRETE - MANBRK TRIMSW SWFAULT FAULT - STAB - L/C HYDR DISCRETE - STAB 28V FAULT

- STAB BRK SW FAULT - STAB 28V FAULT

EXISTING FAULTS? EXISTING

RESET LATCH?

S Q SR Q R

RESET LATCH?

FAULTS?

STAB TRIM

STAB

a

TRIM PANEL a P5 PANEL OVERHEAD P5 OVERHEAD

CSEU +/- 5V DC

CSEU +/- 5V DC

STAB TRIM 28V DC ON OFF

MENU MENU

YES

STAB TRIM 28V DC ON OFF

NO YES

NO

- ALT STAB SW FAULT - ALT STAB SW FAULT - CHNL PIN FAULT - CHNL PINFAULT FAULT - 15V PWR - 15VL/R/C PWR FAULT - FCC 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 - L/R ADC FAULT - L/R ADC FAULT - YSM - YSMBUS BUSXFEED XFEED - YSM - YSMVAL VALDISCRETE DISCRETE

TAT

WARNING WARNING r

STAB TRIMSTAB TRIM

r

CAUTION CAUTION a

a

L/R YSM L/R YSM

EICAS

EICAS

YSM - STABILIZER TRIM - FAULT ANNUNCIATION - STAB TRIM CONT. B767-3S2F Page - 65

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YSM - STABILIZER TRIM- FAULT ANNUNCIATION - STAB TRIM CONT. FCC Fault An FCC fault is set for a FCC bus if a watchdog, parity, coincidence, or engage fault is set for that bus. On the ground, FCC faults have an auto reset when data becomes valid. In the air, FCC faults have an auto reset when the data becomes valid and stays valid for 2.25 seconds. Left, Right, and Center FSPM Fault 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.

Air/Ground 1, 1 and 2, 2 Fault 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 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 continues majority vote. If two air/ground discertes are not valid, the YSM goes to in-air status. The YSM on-ground monitor compares the majority voted status with each of the three air/ground discrete inputs. If there is a 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 majority voting.

- STCM FAULT - MAN TRIM SW FAULT - L/C HYDR DISCRETE STCM FAULT - STAB BRKSW SWFAULT FAULT MAN TRIM L/C HYDR - STAB 28VDISCRETE FAULT - STAB BRK SW FAULT - STAB 28V FAULT

EXISTING FAULTS?

RESET LATCH?

EXISTING FAULTS?

S Q S RQ R

RESET LATCH?

CSEU +/- 5V DC

STAB TRIM STAB

a

PANEL TRIM a P5 OVERHEAD

PANEL P5 OVERHEAD

CSEU +/- 5V DC

STAB TRIM 28V DC MENU MENU

YES YES

ON OFF ON OFF

NO NO

STAB TRIM 28V DC

- ALT STAB SW FAULT -- CHNL PIN SW FAULT ALT STAB FAULT - 15V PWR FAULT CHNL PIN FAULT - FCC L/R/CFAULT FAULT 15V PWR FCC L/R/C FAULT - L/R/C FSPM FAULT L/R/CFAULT FSPM FAULT -- YSM YSM FAULT - AIR/GND 1/1_2/2 FAULT AIR/GND 1/1_2/2 FAULT -- L/R ADC FAULT L/R ADC -- YSM BUSFAULT XFEED -- YSM YSM BUS VAL XFEED DISCRETE

TAT TAT

WARNING

WARNING

- YSM VAL DISCRETE

r

STAB TRIM

r

STAB TRIM

CAUTION

CAUTION

a

a

L/RL/R YSM YSM

EICAS EICAS


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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 FAULT - L/C HYDR DISCRETE - MANBRK TRIMSW SWFAULT FAULT - STAB - L/C HYDR DISCRETE - STAB 28V FAULT

- STAB BRK SW FAULT - STAB 28V FAULT

EXISTING FAULTS? EXISTING

RESET LATCH?

SR Q R

RESET LATCH?

FAULTS?

STAB TRIM

S Q

CSEU +/- 5V DC

STAB

a

TRIM PANEL a P5 PANELOVERHEAD P5 OVERHEAD

CSEU +/- 5V DC

STAB TRIM 28V DC MENU

YES

STAB TRIM 28V DC

ON OFF

ON OFF

MENU

NO YES

NO

- ALT STAB SW FAULT - ALT STAB SW FAULT - CHNL PIN FAULT - CHNL PINFAULT FAULT - 15V PWR - 15VL/R/C PWR FAULT FAULT - FCC - FCC FSPM L/R/C FAULT - L/R/C FAULT - L/R/C FSPM FAULT - YSM FAULT - YSM FAULT - AIR/GND 1/1_2/2 FAULT - AIR/GND 1/1_2/2 FAULT - L/R ADC FAULT - L/R ADC FAULT - YSM - YSMBUS BUSXFEED XFEED - YSM VAL - YSM VALDISCRETE DISCRETE

TAT TAT

WARNING WARNING r r

CAUTION CAUTION a

a

L/R L/R YSMYSM

EICAS EICAS


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STAB TRIM

STAB TRIM

B767-3S2F Page - 70

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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 STAB MOVES OPPOSITE TO CMD DIRECTION

STAB MOTION STOPS

STAB TRIM FUNCTION FAULT CHANNEL CODE FAULT ANY FCC ENGA

0.6 SEC

S

3.0 SEC

R

Q

UNSCHEDULED STAB TRIM UNSCHED STAB TRIM a

ALTERNATE ELECT TRIM

P5 OVERHEAD PANEL

AUTOTRIM MODE CONTROL PROC ARM PROC

LEFT YSM

TAT

AUTOTRIM MODE (RIGHT)

WARNING r

UNSCHD STAB TRIM

CAUTION a

INTERNALS SAME AS LEFT YSM EICAS

RIGHT YSM

YSM - STABILIZER TRIM - FAULT ANNUNCIATION - UNSCHEDULED TRIM B767-3S2F Page - 71

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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

AC MP

EDP

AC MP

AC MP

RIGHT HYD SYS - FLT CONT AC MP

EDP

ADP

L E SLATS RAT

T E FLAPS

STAB TRIM L TAIL SOV

L WING SOV LCCA

STAB TRIM C TAIL SOV

C WING SOV

R WING SOV

LCCA

SPOILER PCA (1-6-12) AIL PCA (LOB-LIB-ROB)

LCCA SPOILER PCA (2-7-11)

SPOILER PCA (3-4-5-8-9-10) 1

AIL PCA (LIB-RIB)

ELEV FEEL CMPTR

AIL PCA (LOB-RIB-ROB)

1 ELEV A/P SERVO

ELEV A/P SERVO

ELEV A/P SERVO



DIR A/P SERVO

DIR A/P SERVO RUDDER PCA RIGHT YAW DAMPER SERVO

DIR A/P SERVO RUDDER PCA

LEFT YAW DAMPER SERVO

RUDDER PCA 1

RUDDER RATIO CHANGER

FLT CONT/HYD - GENERAL DESCRIPTION - FLIGHT CONTROLS Page - 73

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ELEV FEEL CMPTR ELEV PCA (LEFT-RIGHT)

B767-3S2F

R TAIL SOV

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- LOB - LEFT OUTBOARD - LIB - LEFT INBOARD - RIB - RIGHT INBOARD - ROB - RIGHT OUTBOARD

B767-3S2F Page - 74

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AUTOPILOT CONTROL SERVO Purpose The A/P control servo is a hydraulically powered device which converts an electrical signal or command into an output motion or force. This output motion is applied to the aileron power control actuators (PCAs) for aileron control. It also back drives the manual input to move the control wheel. 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.

This causes the detent pistons to clamp the actuator piston to the crank. Detent Trip Monitoring During the synchronizing portion of engagement and subsequently, the FCC compares the autopilot LVDT position with the output LVDT position. If these two do not agree, this is a condition known as Camout. Camout can occur due to manual overriding of the autopilot, due to other autopilot control forces during multichannel operation, because of mechanical jamming or due to a servo malfunction.

SV 2

ELECTRICAL CONNECTOR

EHSV SV 1 FILTER COVER

AUTOPILOT LVDT

RETURN PRESSURE OUTPUT CRANK

SOLENOID VALVE 1

SOLENOID VALVE 2

ELECTRO-HYDRAULIC SERVO VALVE (EHSV)

ELECTRICAL CONNECTOR

HYDRAULIC RETURN PORT PR C

P C1 C2 R

PRESSURE REGULATOR

PRC

OUTPUT LVDT FILTER

P R C

HYDRAULIC PRESSURE PORT ORIFICE (2)

ROLLER

RP

RELIEF VALVE

SERVO PISTON CENTERING SPRING SERVO PISTON OUTPUT LVDT

DISENGAGE SPRINGS INTERNAL CRANK

DETENT PISTON

PIVOT FOR OUTPUT CRANK & INTERNAL CRANK

OUTPUT CRANK TO AILERON CONTROL LINKAGE

AUTOPILOT CONTROL SERVO B767-3S2F Page - 75

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DETENT PISTON

ACTUATOR LVDT

B767-3S2F Page - 76

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LATERAL CENTRAL CONTROL ACTUATORS General Description The lateral central control actuator (LCCA) has a booster section and an 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 booster then moves the wing cables. LCCA Operation - Control Wheel The movement of the booster section piston moves the wing cables. The smaller side of the unbalanced area piston receives a constant hydraulic pressure of 1500 psi from a pressure regulator and relief valve. The control valve also 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 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 piston changes the position of point D which rotates the summing lever about point C and nulls out the input to the control valve. A relief valve and a by-pass check valve allow backdrive of the piston by the other LCCAs. An anti-jam detent allows the input crank to break out (10 lbs control wheel) if the control valve jams and allows operation of the other LCCAs. Hydraulic pressure to the booster section is available when its hydraulic system is pressurized. No electrical signal is necessary. LCCA Operation - Autopilot The autopilot portion of the servo responds to autopilot commands. The servo is enabled by two solenoid valves and controlled by an electrohydraulic servo valve (EHSV). Detent pistons engage the actuator piston to provide servo output. The actuator piston position is detected by the autopilot LVDT. Servo Enabling Two discrete signals are necessary to enable the servo. The first opens solenoid valve 1 and provides hydraulic pressure to solenoid valve 2 and the 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.

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 of the booster crank. Servo Details The control actuator pistons are centered and held open by spring pressure when the autopilot is not engaged. The detent piston is closed by regulated pressure which is equalized between the two sides. The autopilot actuator piston is moved by differential pressure between the two sides as provided from the EHSV. The 26v ac excitation of each LVDT is provided from the associated flight control computer. Autopilot LCCA Operation Before the autopilot is engaged, the LCCA operates as a power booster for manual control and the control section is electrically and hydraulically inactive. 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). The FCC provides commands to the EHSV to synchronize the autopilot actuator piston (actuator LVDT) with the booster actuator piston (output LVDT). When synchronized (1 second) and when other engagement requirements are complete, the FCC provides another dc discrete to engage the detent pistons. The autopilot actuator piston is then clamped to the input crank roller (Point B). LCCA Operation Autopilot commands for aileron deflection are received by the EHSV which moves the autopilot actuator piston. This movement is translated through the input booster crank to both the input arm and the crank. Movement of the booster crank moves the booster actuator and moves of the aileron surface. Hydraulic pressure on the detent pistons is decreased by the pressure regulator 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 can be moved. This causes LVDT disagreement and a detent trip which is detected by the FCC.

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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 automatic control of the airplane in all phases of flight. The AFDS makes stabilizer trim commands when an autopilot is in command.

- Autoflight control systems mode control panel (AFCS MCP) - Flight control computers (FCCs) - Caution and warning annunciators - Autoland status annunciators - Autopilot servos.

Functional Operation 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.

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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 IAS/MACH F/D ON

IAS

VERT SPD

HDG

ALT

L NAV

L B/CRS

A/P ENGAGE

CMD

C

R

CMD

CMD

OFF THR

SEL

5

V NAV

AUTO

OFF

SPD

25 SEL

BANK LIMIT

LOC

DN

HOLD

FL CH

F/D ON

V/S

HOLD

APP

OFF

DISENGAGE

UP

AFCS MODE CONTROL PANEL EADI (2)

ELECTRONIC FLIGHT INSTRUMENT EHSI (2) SYSTEM (EFIS)

AUTOLAND STATUS ANNUNCIATOR (2) AUTOPILOT DISENGAGE SWITCH (2)

A/P DISCONNECT LIGHT A/P CAUTION LIGHT A/P ROLLOUT GUIDANCE SERVO (3)

GO-AROUND SWITCH (2)

A/P LATERAL CONTROL SERVO (3) A/P PITCH CONTROL SERVO (3)

ELECTRICAL POWER FLIGHT MANAGEMENT SYSTEM AIR DATA/INERTIAL REFERENCE SYSTEM RADIO ALTIMETER INSTRUMENT LANDING SYSTEM CENTRAL WARNING SYSTEM ENG IND CREW ALERTING SYSTEM

YSM (2)

FLIGHT CONTROL COMPUTER (3)

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MCDP

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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 damper/stabilizer trim modules (YSMs) and the stabilizer trim control modules (STCMs).

This block diagram shows the AFDS hardware configuration by function. Pilot Control Function The AFCS mode control panel gives the primary interface between the pilots and the AFDS. The thrust management system also uses the mode control panel. Pilot inputs have a direct interface with the flight control computers.

Electronic Flight Instrument System The EADI is the primary display for the AFDS and shows these functions:

I/O, Calculation, and Monitor Functions

- Flight director commands - 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 - Sensor interface and management - Control system interface - Engage and disengage - 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.

Maintenance Function The maintenance control and display panel (MCDP) gives the maintenance function for the AFDS.

DISCRETE WARNING INDICATORS

TMC AFCS MODE CONTROL PANEL

ASA

DISCRETE CAUTION INDICATORS

EADI EHSI STAB POSN

FMC EFIS DUAL REDUNDANT NAVIGATION SENSORS

ELEV SERVOS

TO SURFACE PCA(S)

AILERON SERVOS

TO SURFACE PCA(S)

DIRECT SERVOS

TO SURFACE PCA(S)

FLAP POSN

SPD BRAKE HDLE POSN FLIGHT CONTROL COMPUTER SLAT SW LOGIC

ADIRU

AUTOPILOT SERVOS

YSM

MCDP

HYD VALID LOGIC

STCM AUTO STAB TRIM

ILS DISENGAGE SWITCHES

MAINTENANCE FUNCTION DFDAU

AIR/GND LOGIC

RA

AIRPLANE CONFIG SENSORS

TRIPLE REDUNDANT NAVIGATION SENSORS

F/D SOURCE SEL SW

HDG REF SWITCH

G/A SWITCHES MISCELLANEOUS PILOT INPUTS

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FLIGHT RECORDER FUNCTION

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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 - A/P DISC LIGHT - A/P CAUTION LIGHT

FLAP LEVER - LEADING EDGE SLAT SWITCH P55 GLARESHIELD PANEL - MODE CONTROL PANEL

THRUST LEVERS - GO-AROUND SWITCHES

P3 F/O INSTRUMENT PANEL - EADI - INSTRUMENT SOURCE SELECT SWITCH PANEL - ASA - HEADING REFERENCE SWITCH P61 AUXILIARY PANEL - AIRBORNE DATA LOADER

YAW DAMPER SERVOS P1-1 CAPT INSTRUMENT PANEL - EADI - INSTRUMENT SOURCE SELECT SWITCH - ASA

CONTROL WHEELS - A/P DISCONNECT SWITCHES

ELEV A/P SERVOS

LATERAL CENTRAL CONTROL ACTUATORS

L FCC (E1-3) C FCC (E1-4) R FCC (E1-5)

FWD

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DIRECTIONAL A/P SERVOS


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MODE CONTROL PANEL - FUNCTIONS Switches, Lights and Controls Autothrottle Arm Sw- Autothrottle is Armed by use of this toggle switch when in A/T ARM position. This switch provides DC PWR to the autothrottle servo. OFF disables autothrottle Autothrottle Mode Select switches. The FMA displays autothrottle mode of operation. The following are Autothrottle modes: THR -When the THR light bar is on engine(s) are at MAX EPR or MAX Thrust as selected from the TMSP. 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. VNAV- When the VNAV light bar is on engine(s) Limits, Speed, throttle position, pitch of the aircraft and performance are being determined by the FMC. 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 TMSP. Lateral Navigation-LNAV light bar is when LNAV is either ARMED or Engaged as displayed on the FMA. When L NAV is operating the aircraft roll and heading is being controlled by the FMC. 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. AIRSPEED / MACH CONTROL Switch- changes the speed display numbers and transfers the command to and from the FMC to the MCP when in performing a speed intervention. AIRSPEED / MACH SEL- SEL changes the display between airspeed and mach. Mach must be above .40 to display mach. Flight Director switches- F/D controls the FD displays on the left and right PFD ADI respectively.

HEADING DISPLAY- This HEADING DISPLAY is capable of displaying 000 to 359 degrees in increments of one degree. Initialized at 000 on power up. 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. Depressing SEL causes the aircraft to obtain the heading selected on the window. Another way is to press select then turn the heading to the desired heading and the A/P will follow as the knob is turned. HEADING HOLD SWITCH- This flowbar illuminated indicates HDG has been held. VERTIICAL SPEED DISPLAY- The VERTIICAL SPEED DISPLAY window displays Vertical Speed (V/S) from +6000 to -8000 ft/min in 100 ft/min increments. VERTICAL SPEED RATE WHEEL-dials in a selected V/S. VETTICAL SPEED SLECET SWITCH-The V/S switch activates the V/S mode. ALTITUDE DISPLAY WINDOW- ALT shows from 0-50,000 feet in 100 feet increments. ALTITUDE SELECT KNOB/SWITCH-Turning of the knob clockwise increases the ALT in the ALTITUDE DISPLAY WINDOW. ALTITUDE HOLD-HOLD This flowbar when pressed engages ALT hold. AUTOPILOT DISENGAGE BAR-DISENGAGE is a switch that removes power from the AP servos by applying a ground to the servos. 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 APP has been selected it is possible to have multiple CMD flow bars illuminated. These AP(s) do not engage until less than 1500 feet RA. BACK COURSE SELECT- This flowbar illuminated indicates Back course selected LOCALIZER SELECT- This flowbar illuminated indicates LOC selected. APPROACH SELECT- This flowbar illuminated indicates APP selected.

AUTOTHROTTLE ARM SWITCH

SPEED DISPLAY MACH / IAS

THRUST MODE SELECT SWITCH

FLIGHT LEVEL CHANGE MODE SELECT SWITCH/LIGHT

SPEED MODE SELECT SWITCH

AIRSPEED / MACH SELECT SWITCH

LATERAL NAVIGATION MODE SELECT SWITCH/LIGHT

AIRSPEED/MACH CONTROL AND LOCAL/REMOTE SPEED SELECT SWITCH

VERTICAL SPEED DISPLAY

HEADING DISPLAY HEADING SELECT ANK ANGLE LIMIT SELECTOR

VERTICAL NAVIGATION MODE SELECT SWITCH/LIGHT

ALTITUDE DISPLAY

VERTICAL SPEED RATE UP/DOWN SELECT WHEEL VERTICAL SPEED SELECT SWITCH/LIGHT

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APPROACH SELECT SWITCH/LIGHT ALTITUDE SELECT SWITCH

HEADING HOLD SELECT SWITCH/LIGHT

MODE CONTROL PANEL - FUNCTIONS ATA 22-00

AUTOPILOT ENGAGE SWITCHES/LIGHTS

LOCALIZER SELECT SWITCH/LIGHT

ALTITUDE HOLD SWITCH/LIGHT

B767-3S2F

BACK COURSE SELECT SWITCH/LIGHT


FLIGHT DIRECTOR ON/OFF SWITCH (2)

AUTOPILOT DISENGAGE SWITCH/LEVER

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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

WARNING

IAS/MACH F/D ON

CAUTION

THR

CAPTAIN AND F/O MASTER WARNING AND CAUTION LIGHTS (P7)

IAS MACH

OFF

VERT SPD

HDG

ALT

L

L NAV

SEL

5

V NAV

C

R

CMD

CMD

F/D ON

25

AUTO

OFF

SPD

CMD

B/CRS

A/P ENGAGE

SEL

LOC

DN

BANK LIMIT

OFF

HOLD

FL CH

HOLD

V/S

APP

DISENGAGE

UP

MODE CONTROL PANEL (P55) STATUS PITCH MODES OPERATE

G/S

LOC

FLARE

CMD

ROLLOUT

MODE ANNUNCIATION (TYP)

CMD

AUTOLAND STATUS

F/D

TO ALT HOLD ALT CAPT

EADI PITCH MODES ROLL MODES ARMED ARMED

OPERATE

LOC

TO

SPD

FLARE

ROLLOUT

HDG HOLD

G/S

VNAV

B/CRS

HDG SEL

LNAV

LNAV LOC

VNAV SPD

TRK

252

M

AUTOLAND STATUS ANNUNCIATION (P1-3)

AUTOPILOT WARNING ANNUNCIATION (RED)(P1-3)

RED TAT +12C

AUTOPILOT CAUTION ANNUNICATION (AMBER)(P1-3)

AUTOPILOT DISC AUTOPILOT

A/P DISC

AUTO PILOT

ROLLOUT TUKWA

ATT

ONP

FLAP LIM

TO

YELLOW

GA

VNAV ALT

EICAS UPPER DISPLAY UNIT KEY:

HEADING CURSOR

BENSN

MODES

ENGAGED

ARMED

LIMIT

STATUS

GREEN

N/A

N/A

PITCH

GREEN

WHITE

GREEN

ROLL

GREEN

WHITE

N/A

20

EHSI

AFDS - ANNUNCIATION AND WARNING - GENERAL DESCRIPTION B767-3S2F Page - 89

LAND 2 (GREEN) LAND 3 (GREEN)

B/CRS

VNAV PTH SPD LIM

TEST 2

G/S

GA

P/RST

TST

V/S

FLARE

1

NO LAND 3 (AMBER) NO AUTOLAND (AMBER)

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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 multichannel 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.

(A)

LAND 3 (GREEN)

(B)

LAND 2 (GREEN)

RESET SWITCH TEST SWITCH (A FACES)

AUTOLAND STATUS

AUTOLAND STATUS 1

P/RST

1 P/RST

TEST

TEST

(A)

NO LAND 3 (AMBER)

(B)

NO AUTOLAND (AMBER)

2

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2

TEST SWITCH (B FACES)

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AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE -1 General 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.

Autopilot Channel Selection Autoland status annunciations are different with the number of autopilot channels engaged. Displays and related operations are described for single, dual, and triple channel selection.

DEGRADATION DISPLAYS

AUTOLAND STATUS

CRUISE CLIMB

1

IN AIR, CLIMB, OR CRUISE

P/RST TEST

- MONITOR OF FCC, SENSORS AND INTERFACES - NO AUTORESET

IN AIR ON GROUND - ON GROUND MONITOR DYNAMIC TEST OF SENSORS AND INTERFACES SINGLE FAULT

SINGLE FAULT

2

FAULT CONDITION BEFORE APPROACH SELECT

NO LAND 3

- MANUAL RESET PUSH AND HOLD REINTIALIZES VOTING CIRCUIT FOR REASSESSMENT OF CONDITION

NO LAND 3

- AUTORESET WHEN CONDITION CLEARED - DISPLAY REMOVED DURING RESET - DISPLAY REMAINS BLANK WHEN RESET RELEASED IF FAULT CONDITION CLEARED

- DISPLAY REMOVED

MULTIPLE FAULTS

MULTIPLE FAULTS

NO AUTOLAND

NO LAND 3

- AUTORESET WHEN APPROACH SWITCH PUSHED - VOTING CIRCUIT REINITIALIZED FOR REASSESSMENT OF CONDITION

- DISPLAY REMOVED IF FAULT CONDITION CLEARED BEFORE APPROACH SELECTED MULTIPLE FAULTS

NO AUTOLAND

- AUTORESET WHEN CONDITION CLEARED

- MANUAL RESET

- DISPLAY REMOVED

- OPERATION SAME AS FOR NO LAND 3 ABOVE

NO AUTOLAND

- AUTORESET WHEN APPROACH SWITCH PUSHED

- OPERATION SAME AS FOR NO LAND 3 ABOVE

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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 FAULT CONDITION AFTER APPROACH SELECT SINGLE FAULT

APPROACH SELECT BELOW 1500 FT TRIPLE CHANNEL TRIPLE ENGAGE CHANNEL BELOW ARM 600 FT

AUTOLAND STATUS 1 P/RST TEST 2

1 BELOW ALERT HEIGHT

NO LAND 3

- MANUAL RESET BELOW 1500 FT

BELOW ALERT HEIGHT

LAND AND FCC DISENGAGE

LAND 3

LAND 3

NO LAND 3

- DISPLAY REMOVED AND LATCHED OFF UNTIL LANDING AND DISENGAGE - FAULT CONDITION LATCHED IN FCC - NO UPGRADE FOR CLEARED FAULT CONDITION

SINGLE FAULT LAND 2

LAND 3

NO LAND 3

- MANUAL RESET

MULTIPLE FAULTS

- MANUAL RESET REVERTS TO ON GROUND MONITOR

SINGLE FAULT

MULTIPLE FAULTS NO AUTOLND

LAND 2 NO AUTOLND

NO AUTOLND

- DISPLAY LATCHED ON UNTIL LANDING AND DISENGAGE - DISPLAY REMOVED DURING MANUAL RESET WHILE SWITCH IS HELD - FAULT CONDITION LATCHED IN FCC - NO UPGRADE FOR CLEARED FAULT CONDITION

- NO LAND 3 DISPLAY REMOVED WITH CONDITION EXISTING MULTIPLE FAULTS

- NOT REMOVED BY MANUAL RESET

NO AUTOLND

1

IF MULTICHANNEL IS NOT ENGAGED BY 600 FT, THE ASA SHOWS NO AUTOLAND. NO UPGRADE FOR CLEARED FAULT CONDITION. ASA AUTORESETS WHEN A/P IS DISENGAGED OR NEW MODE IS SELECTED.


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- MANUAL RESET REVERTS TO ON GROUND MONITOR

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AFDC - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE - 3 General This page continues the autoland status annunciator diaplay sequence.

CRUISE

APPROACH SELECT AND DUAL CHANNEL ARM

AUTOLAND STATUS 1 P/RST TEST

BELOW 1500 FT AND DUAL CHANNEL ENGAGE

APPROACH SELECT - DUAL CHANNEL ARM - MONITOR OF FCC, SENSORS AND INTERFACES

2

BELOW 600 FT BELOW ALERT HEIGHT

SINGLE FAULT CONDITION AFTER APPROACH SELECT AND DUAL CHANNEL ARM

STATUS AND DEGRADATION DISPLAYS

SINGLE FAULT BELOW 1500 FT AND DUAL CHANNEL ENGAGE NO LAND 3

- ASA REMAINS BLANK - FAULT CONDITION LATCH IN FCC(S)


BELOW ALERT HEIGHT

LAND 2

LAND 2

NO LAND 3

NO LAND 3

SINGLE FAULT

NO LAND 3

SINGLE FAULT

- RESET REVERTS TO ON GROUND MONITOR

MULTIPLE FAULTS

NO AUTOLND

- DISPLAY LATCHED ON UNTIL LANDING AND DISENGAGE - DISPLAY REMOVED DURING MANUAL RESET WHILE SWITCH IS HELD - FAULT CONDITION LATCHED IN FCC(S) - NO UPGRADING FOR CLEARED FAULT CONDITION

LAND 2

LAND 2

NO LAND 3

NO LAND 3

- FAULT CONDITION LATCH IN FCC(S) MULTIPLE FAULTS

- FAULT CONDITION LATCH IN FCC(S)

1

MULTIPLE FAULTS

NO AUTOLND


1

1

NO AUTOLND

NO AUTOLAND

- NOT REMOVED BY MANUAL RESET - IMMEDIATE NO AUTOLAND

NO LAND 3 SHOWS AFTER LANDING AND DISENGAGE.

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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 ALERT HEIGHT


NO LAND 3

SINGLE FAULT


NO AUTOLND


- FAULT CONDITION LATCHED IN FCC(S) 1 MULTIPLE FAULTS

NO AUTOLND

- NOT REMOVED BY MANUAL RESET NO LAND 3 SHOWS AFTER LANDING AND DISENGAGE AFDS - FUNCTIONAL DESCRIPTION - ASA - DISPLAY SEQUENCE - 4 1

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MODE CONTROL PANEL INTERNALS General To accomplish the requirement that a single failure will not cause total loss of the AFDS mode control panel function, the mode control panel contains two separate control channels with a common display. Each channel is powered 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.

Monitors Power supply monitoring hardware supplies appropriate reset signals to the 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 hardware watchdog timer. If the timer is not restarted every 150 to 200 milliseconds, the critical portions of the I/O system and associated CPU are disabled allowing the alternate CPU to assume control without interference.

MODE CONTROL PANEL INTERNALS B767-3S2F Page - 101

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ENGAGE LOGIC - SIMPLIFIED General 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 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 engagement and rudder engagement. Elevator/Aileron Engagement Arm is initiated by command request from the MCP if single channel engaged and by command request and approach mode with another channel engaged. With the arm latch set, the elevator/aileron arm relay is energized which will 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.

Elevator and Aileron Engage The engage function is initiated when the arm state is achieved. The elevator/ aileron engage relay operates similarly to the arm relay to energize the engage solenoid in the elevator and lateral autopilot servos. Rudder Arm Rudder arm is initiated by multichannel operation which is enabled only in the approach mode. The rudder arm solenoid is energized through the rudder arm relay similar to elevator/aileron arm. Rudder Engage Upon achieving rudder arm status, the rudder engage logic is initiated leading to the energizing of the rudder engage solenoid when multichannel engaged.

28V DC SERVO

ELEV ARM

ELEVATOR ARM SOLENOID

AIL ARM ARM & ENGA DISC 429 XMTR

ELEV & AIL ARM LOGIC

429 RCVR

ELEV SERVO ENGAGE

DATA MGMT 429 XMTR

429 RCVR

D A T A

CMD

ELEV/AIL ENGA RLY

AFCS MCP

RUDDER ARM RUDDER ARM RLY

RUDDER SERVO ENGAGE RUDDER ENGAGE LOGIC

429 RCVR

FLIGHT CONTROL COMPUTER (TYP)

ENGAGE LOGIC - SIMPLIFIED Page - 103

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A/P PITCH CONTROL SERVO AILERON ARM SOLENOID AILERON ENGAGE SOLENOID A/P LATERAL CONTROL SERVO

RUDDER ARM LOGIC

RUDDER ENGA RLY X-CHANNEL BUSSES

AIL SERVO ENGAGE

ELEV & AIL ENGAGE LOGIC

M A N A G E M E N T

OFF DISENG BAR

B767-3S2F

ELEVATOR ENGAGE SOLENOID

ELEV/AIL ARM RLY

RUDDER ARM SOLENOID RUDDER ENGAGE SOLENOID

A/P ROLLOUT GUIDANCE SERVO

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ENGAGE LOGIC DETAILS - HARDWARE MONITORS General The hardware monitor logic receives the arm or engage logic developed previously and if conditions are appropriate, passes this logic to the arm and engage and for rudder arm and engage. 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.

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 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 until touchdown and disconnect. The ASA will then annunciate NO LAND 3. If the autopilot is triple channel engaged a dual or triple channel camout will cause an autopilot disconnect annunciation and NO AUTOLAND on the ASA.

WHEEL DISENGAGE (FROM ELEV/AIL ARM LOGIC)

ELEV/AIL ARM

COMPUTER INHIBIT

FROM OTHER CHANNELS

LEFT SERVO ENGA RIGHT DETENT

1

DC

ELEV/AIL ENGA RELAY

A

WHEEL DISENGAGE ELEV/AIL ENGAGE

(FROM ELEV/AIL ENGA LOGIC) LEFT DETENT

ELEV/AIL ARM RELAY

PRE ARM

COMPUTER VALID

FROM COMPUTER MONITOR

DC POWER SUPPLY VALID

POWER SUPPLY VALID

LEFT CAMOUT

PRE ENGA

CAMOUT LOGIC B

RIGHT CAMOUT

RIGHT SERVO ENGA

2

AIL DETENT

FOREIGN SERVO ENGA

AIL SERVO DETENT COMPARATOR

AIL SURFACE LVDT POSITION

ELEV SERVO ELEV SURFACE RUD SERVO

DETENT LOGIC LOW IF LVDT(S) DIFFER BY >3` (AIL & RUD) >4` (ELEV)

MY CAMOUT

MY DETENT TO OTHER CHANNELS

ELEV DETENT TO AFDS ANNUNCIATIONS

DC

RUD DETENT

RUD SURFACE 2 IN COMMAND 3 IN COMMAND TWO SAME INNER LOOP FAILURE

(FROM RUDDER ARM LOGIC) RUDDER ARM

A

NON-ISO DISCONNECT

DC B POWER SUPPLY VALID RUDDER ENGAGE (FROM RUDDER ENGA LOGIC) FOREIGN SEVO ENGA

TWO SAME AUTOLAND SENSORS FAILURE FLIGHT CONTROL COMPUTER (TYP) 1

2

INHIBIT LOGIC LOW DURING POWER UP TEST ONLY

USED DURING DUAL CHANNEL OPERATION ONLY - IF FOREIGN CHANNEL CAMOUT YOU GET A DETENT TRIP (DETENT LOGIC LOW)

ENGAGE LOGIC DETAILS - HARDWARE MONITORS B767-3S2F Page - 105

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RUDDER ARM RELAY

RUDDER ENGA RELAY

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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 DETENT TRIP

RELATIVE LEFT

SERVO ENGA DETENT TRIP X CHANNEL LEFT

RELATIVE RIGHT

1 X CHANNEL RIGHT MY

SERVO ENGA DETENT TRIP

LEFT FCC

1 SAME AS LEFT FCC

1

CROSS-CHANNEL DATA CENTER FCC

AILERON COMMAND ELEVATOR COMMAND RUDDER COMMAND

1

DISCRETE WORDS GENERAL PURPOSE WORDS SAME AS LEFT FCC

IRS SIGNALS ADC SIGNALS ILS SIGNALS FMC SIGNALS RADIO HEIGHT RIGHT FCC

FLAP POSITION STAB POSITION

AFDS GENERAL - INTERFACE - FCC CROSS-CHANNEL DATA B767-3S2F Page - 107

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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 BUS LEFT

28V DC (BAT)

AUTO FLT WARN DC

5V AC 28V DC PWR

MODE CONT PANEL L

115V MAIN AC BUS LEFT 115V AC BUS CENTER 28V DC BUS CENTER

AFCS WARN PWR

CAPT ASA

SERVO ARM PWR

L FCC SERVO

SERVO ENG PWR

L LOGIC POWER

PRIMARY PWR L FCC

L FCC PWR

L SERVO POWER

C FCC PWR SAME AS L FCC

C FCC SERVO

C FCC 28V DC BUS RIGHT

C SERVO POWER

MODE CONT PANEL SAME AS L FCC

R FCC SERVO 115V MAIN AC BUS RIGHT

R FCC PWR

R FCC

5V AC

R SERVO POWER

SAME AS CAPT ASA

5V AC MASTER DIM MASTER TEST MASTER DIM MASTER TEST

LIGHT PLATE PWR IND PWR L GND FOR TEST IND PWR R GND FOR TEST

F/O ASA R LOGIC POWER AFCS MCP

AFDS POWER DISTRIBUTION B767-3S2F Page - 109

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AFDS FCC POWER DISTRIBUTION FCC AC Power Each autopilot servo receives 26vac for LVDT excitation from its FCC.

26V AC

LVDT EXCITATION L FCC

SERVO POSN LVDT SURF POSN LVDT L LCCA SERVO POSN LVDT SURF POSN LVDT L ELEVATOR A/P SERVO SERVO POSN LVDT SURF POSN LVDT L DIRECTIONAL A/P SERVO

SAME AS L FCC

26V AC

SAME AS L COMPONENT C LCCA SAME AS L COMPONENT C ELEVATOR A/P SERVO

C FCC

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

26V AC

SAME AS L COMPONENT R LCCA SAME AS L COMPONENT R ELEVATOR A/P SERVO SAME AS L COMPONENT R DIRECTIONAL A/P SERVO

AFDS FCC POWER DISTRIBUTION B767-3S2F Page - 111

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AUTOLAND POWER SWITCHING - FUNCTIONAL DESCRIPTION

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 from one source.

General

Operation

It is necessary for the airplane to have Category III B capability. The categories are established by the regulatory agencies based on these factors:

The standby electrical system is an independent power source to supply the center autopilot during autoland operation. During autoland operation, the battery charger (which operates in the T-R mode as a constant voltage source) is the third power source. The battery provides instantaneous backup if the battery charger fails.

- Airplane certification - Airport facilities - Flight crew proficiency - Airline maintenance. The runway visual range (RVR) and decision height (DH) requirements established for category III B are typically 150 - 700 ft RVR and 0 ft DH. These numbers are based on a fail-operational capability. This is the level of redundancy necessary such that if there is a single failure that occurrs below the alert height, the landing may be continued with the remainder of the automatic system. For a fail-operational capabilty, triple redundancy of these is necessary: - Power sources - Engaged FCCs - Sensors - Servos. 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

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 bus power control unit (BPCU). The BPCU sends an autoland lockout command to the left and right generator control units to inhibit BTB close command logic. At least one BTB must be open and one GCB closed to energize the autoland relays (K526 and K527). For backup protection, the circuit to the BTB close coil is open to interrupt any BTB close command when the autoland relay is energized. When either K526 or K527 is energized, if the static inverter is operational, relay K2127 energizes. This lets center isolation relay, K123, energized. If the battery switch is ON, K2128 and K107 (center bus transfer relay) energize switching the center buses. Additional contacts of relay K2128 bypass K2127 after isolation occurs. After center bus transfer is complete, the bus isolated ground input (from K108) is removed when K107 is energized.

APC

APU GEN

INV PWR TRU T-R

BAT. BUS CONT

B

STBY POWER SW

K106 MAIN BAT. XFR RLY

AUTO

C BUS DC AUTOLAND

CAT 3 BUS ISO

TO K137 IRS DC PWR DISC

LEFT

HOT BAT BUS

AUTOLAND COMMAND TO BUS POWER CONTROL UNIT

CAT 3 BUS ISO

CAT 3 BUS ISO RIGHT

K122 ISOL REQUEST RELAY

AUTOLAND BUS ISOL

2

L FCC

BAT. XFR CONT

K527 RIGHT AUTOLAND RLY

C FCC AUTOLAND BUS ISOL R BTB

R GEN

R GCB

1

WHEN ENERGIZED CONTACTS NOT SHOWN PREVENT L BTB FROM TRANSFERRING

2

WHEN ENERGIZED CONTACTS NOT SHOWN PREVENT R BTB FROM TRANSFERRING

K123 CENTER ISOL RLY

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BUS ISOLATED

BATTERY SWITCH

L FCC

ON

BUS ISOLATED

DC TIE CONTROL UNIT

T-R

R MAIN LOW F/O INST BUS VOLTAGE SENS


C FCC BUS ISOLATED R FCC

C

AUTOLAND POWER SWITCHING - FUNCTIONAL DESCRIPTION B767-3S2F

K107 CENTER BUS XFR RELAY

28V K108 DC TIE RELAY

AUTOLAND BUS ISOL

R FCC

C BUS CONT

K104 MAIN BAT. RLY

MAIN AC BUS - R

C

DC BUS - R

B

C BUS DC POWER

K526 LEFT AUTOLAND RELAY

BAT.

C BUS AC AUTOLAND

TO AC STBY RLY

BAT.

OFF

EPC EXTERNAL POWER

A

STANDBY INVERTER

115V AC BUS - C

1

C BUS AC POWER

28V DC BUS - C

STBY BUS L GCB

DC BUS - L

L BTB

MAIN AC BUS - L

L GEN

A

L MAIN LOW CAPT INST BUS VOLTAGE SENS

B767-3S2F Page - 114

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AUTOLAND POWER SWITCHING - FUNCTIONAL DESCRIPTION CONT. Bus Isolation Signal The bus isolation sigmal shows that the buses are isolated. All three channels 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.

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 directly sense the GCB position. An open GCB causes the BPCU to remove the autoland lockout command from the corresponding GCU and allows 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 of failure.

APC

APU GEN

INV PWR TRU T-R

BAT. BUS CONT

B

STBY POWER SW

K106 MAIN BAT. XFR RLY

AUTO

C BUS DC AUTOLAND

CAT 3 BUS ISO

TO K137 IRS DC PWR DISC

LEFT

HOT BAT BUS

AUTOLAND COMMAND TO BUS POWER CONTROL UNIT

CAT 3 BUS ISO

CAT 3 BUS ISO RIGHT

K122 ISOL REQUEST RELAY

AUTOLAND BUS ISOL

2

L FCC

BAT. XFR CONT

K527 RIGHT AUTOLAND RLY

C FCC AUTOLAND BUS ISOL R BTB

R GEN

R GCB

1

WHEN ENERGIZED CONTACTS NOT SHOWN PREVENT L BTB FROM TRANSFERRING

2

WHEN ENERGIZED CONTACTS NOT SHOWN PREVENT R BTB FROM TRANSFERRING

K123 CENTER ISOL RLY

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BUS ISOLATED

BATTERY SWITCH

L FCC

ON

BUS ISOLATED

DC TIE CONTROL UNIT

T-R

R MAIN LOW F/O INST BUS VOLTAGE SENS


C FCC BUS ISOLATED R FCC

C

AUTOLAND POWER SWITCHING - FUNCTIONAL DESCRIPTION CONT. B767-3S2F

K107 CENTER BUS XFR RELAY

28V K108 DC TIE RELAY

AUTOLAND BUS ISOL

R FCC

C BUS CONT

K104 MAIN BAT. RLY

MAIN AC BUS - R

C

DC BUS - R

B

C BUS DC POWER

K526 LEFT AUTOLAND RELAY

BAT.

C BUS AC AUTOLAND

TO AC STBY RLY

BAT.

OFF

EPC EXTERNAL POWER

A

STANDBY INVERTER

115V AC BUS - C

1

C BUS AC POWER

28V DC BUS - C

STBY BUS L GCB

DC BUS - L

L BTB

MAIN AC BUS - L

L GEN

A

L MAIN LOW CAPT INST BUS VOLTAGE SENS

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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 NO AUTOLAND

28V A

R NO LAND 3 OR NO AUTOLAND

D A T A

MY SVO ENGA MY SVO ARM L SVO ENGA

TRIPLE ARM

L SVO ARM

S R

OS

Q

R SVO ENGA R SVO ARM APPROACH ARM BELOW ALERT HEIGHT (<200' RA) A

MY DISCONNECT L DISCONNECT

GA INITIATE

R DISCONNECT

4 SEC BUS ISOLATED

SOFTWARE

TO NO LAND 3 LOGIC

HARDWARE LEFT FLIGHT CONTROL COMPUTER

AFDS GENERAL - POWER ISOLATION LOGIC TRAINING MANUAL ATA 22-00

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K122 ISOLATION REQUEST RELAY TO R FCC TO C FCC

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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.

B-767 Autoland and Go Around

1500’ (Radio Altimeter) * ASA: LND 3 or LAND 2 * Remaining Autopilots Engage * FLARE & ROLLOUT arm * Rudder active for asymmetrical thrust

* When APPROACH MODE is armed and Autopilot is engaged, remaining operational ones arm. 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).

500’ RA * RUNWAY ALIGNMENT (not annunciated)

Above 400’ RA * Engage any Pitch / Roll Mode back to original Autopilot)

* Autothrottle reduces to old 2000 FPM rate of climb.

200’ RA ASA: NO LAND 3 inhibited * Center buses Auo Transfer inhibited

330’ RA * AUTO NOSE UP TRIM (not annunciated)

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.

Below 400’ RA * To disengage GA disengage Autopilots and both Flight Directors.

45’ RA * FLARE annunciates * Power goes to Go-Around thrust * Pitches to hold existing speed, “Bug” speed of ALPHA speed, whichever is higher. * Roll maintains ground track that exists at the time of engagement.

25’ RA * IDLE annunciates AUTO GO-AROUND * 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) * Autopilot holds nse down yoke * FLARE disengages * Autobrakes should engage

AUTOLAND SEQUENCE B767-3S2F Page - 119

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* Selecting Reverse disengages Autothrottle (no warnings) * Reversers deploy Auto Speed Brakes (if not already deployed) * Disconnect Autopilot to turnoff runway

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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 SERVO DRIVE

EICAS AND EFIS DISPLAYS

AIRPLANE & NAVIGATION SENSORS

TMSP CONTROLS


MCP CONTROLS AND DISPLAYS AFCS MODE CONTROL PNL

THRUST MANAGEMENT COMPUTER

THRUST MANAGEMENT SYSTEM - INTRODUCTION B767-3S2F Page - 121

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ELECTRONIC ENGINE CONTROLS AND SENSORS

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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 CONTROL PANEL (P55)

P3 INSTRUMENT PANEL - THRUST MODE SELECT PANEL

LE SLAT SWITCH (BEHIND PANEL)

P1 INSTRUMENT PNL - A/T DISCONNECT LIGHT SERVO MOTOR (UNDER FLOOR) MICRO SWITCH PACK ASSEMBLY (UNDER FLOOR)

E1-3 SHELF - THRUST MGMT COMPUTER GO-AROUND SWITCH (2)

AUTOTHROTTLE DISENGAGE SWITCH (2)

TMS - COMPONENT LOCATION B767-3S2F Page - 123

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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 ANGLE RESOLVER

AUTOTHROTTLE SERVOMOTOR

TLA RESOLVER LINKS (2) THRUST LEVER ANGLE RESOLVER

TMS - THROTTLE COMPONENT LOCATION B767-3S2F Page - 125

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TMS - THRUST MODE SELECT PANEL General Description 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.

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. The maximum derate is 25%.

SET DERATE SELECT SWITCHES

THRUST LIMIT MODE SELECT SWITCHES (4)

TEMPERATURE SELECTOR

TMS - THRUST MODE SELECT PANEL B767-3S2F Page - 127

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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 B767-3S2F Page - 129

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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 TEMPERATURE

REFERENCE LIMIT DISPLAY SELECTED TEMPERATURE

MAXIMUM LIMIT MARKER

TAT _________ + 12c

REFERENCE THRUST MODE

+10c

REFERENCE/ TARGET CURSOR

TO

107.1

20.0

10 6

107.1

20.0

10

2

THRUST DISPLAY

2

6

N

1

385

THRUST POINTER 385 PRIMARY THRUST PARAMETER

EGT

EICAS DISPLAY

DISPLAY

COMPUTER

ENGINE STATUS

EVENT RECORD

L

AUTO

BRT

THRUST REF SET L

R

BOTH

PULL

R

MAX IND RESET

MANUAL THRUST SET KNOB

EICAS CONTROL PANEL

TMS - OPERATION - ENGINE EICAS DISPLAY - THRUST LIMIT FUNCTIONS B767-3S2F Page - 131

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TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK DIAGRAM Thrust Control The autothrottle servo and clutch move the throttle system. The thrust lever resolvers electrically transmit the throttle lever angle (TLA) to each engine. Dual electronic engine controls (EECs) control fuel flow and send engine data back to the TMC. EICAS Display The EICAS display shows N1 data and thrust limit modes. EADI The autothrottle mode shows on the EADI. A/T DISC Light The A/T DISC light is an amber caution light controlled by the TMC internal monitors and self-test functions. Sensor Inputs These are the sensor inputs: - Attitude and rate from the IRS - Air data from the ADC - Airplane configuration discretes. Digital Signal Inputs Digital inputs are from these units: - MCP - Mode selection - EEC - N1, TLA, and EEC status data - FMC - Thrust limits from the MCDU, mode data, and performance data - ADC - Air data - IRU - Inertial reference data - MCDP - Test data.

Digital Signal Outputs Digital outputs go to these units: - MCP - TMS mode status and vertical speed - EEC - Engine trim - FMC - TMS mode status, engine bleed status, N1, flap position, and temperature selection - MCDP - Fault, ground test, and configuration data. Analog Input Signals Analog inputs are from these units: - EEC discrete GE card - A/T disconnect and GA switches - Air ground relays - Flap slat position module - Leading edge slat switch. MCP The MCP gives this data: - A/T ARM - A/T mode selection - IAS/mach selection - Elevator speed commands. TMSP The TMSP gives thrust limit mode selection and assumed temperature input. FMC The FMC gives commands to the A/T in VNAV mode. These are the inputs from the FMC: - Thrust limit modes - IAS and mach targets - N1 - GMT - A/T mode requests.


L NAV

IAS OFF SEL

THR

V NAV

B

OFF

SPD

FL CH

R ENGINE EEC DATA

AUTOTHROTTLE FUNCT

C

MCP

TO GA

CLB

CON

CRZ

1

2

DUAL CHANNEL ELECTRONIC ENGINE CONTROL

TO R ENGINE

NI REF

FMC

EEC TRIM

SERVO THRUST LIMIT FUNCT

BLEED CONFIG CARDS

CLUTCH RESOLVER SYNCHROS

A

FUEL METERING

L-PIMU

TEMP SEL

POWER

LEFT ENGINE

TMSP INPUT SIGNAL MANAGEMENT

R

DFDAU

C L

ADIRU(S)

EEC TRIM CONTROL

A/G RELAY

INTERNAL MONITOR AND SELF-TEST

L FSPM LE SLAT SW 28V DC STBY BUS 115V AC L MAIN BUS 28V DC L MAIN BUS

EADI

EFIS SYMBOL GENERATORS

B C

EICAS COMP

A/T DISC

a

TAT + 12C

AFCS WARN

+38C

COLLINS 96.1

A

12

GRD TEST SEL UP

TMC AC

POWER SUPPLY

DOWN

YES/ADV

ON/OFF

FLT FAULTS

GND TEST

8

26.1

4

8 N1

318

320

TMC SERVO

28V DC L MAIN BUS

TMC

MCDP

TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK DIAGRAM B767-3S2F Page - 133

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115.2 4

NO/SKIP

EGT

TMC P11 OVERHEAD CB PANEL

96.1

D-TO

54.9

EICAS DISPLAY UNIT

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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.


L NAV

IAS OFF SEL

THR

V NAV

B

OFF

SPD

FL CH

R ENGINE EEC DATA

AUTOTHROTTLE FUNCT

C

MCP

TO GA

CLB

CON

CRZ

1

2

DUAL CHANNEL ELECTRONIC ENGINE CONTROL

TO R ENGINE

NI REF

FMC

EEC TRIM

SERVO THRUST LIMIT FUNCT

BLEED CONFIG CARDS

CLUTCH RESOLVER SYNCHROS

A

FUEL METERING

L-PIMU

TEMP SEL

POWER

LEFT ENGINE

TMSP INPUT SIGNAL MANAGEMENT

R

DFDAU

C L

B

ADIRU(S)

EEC TRIM CONTROL

A/G RELAY

INTERNAL MONITOR AND SELF-TEST

L FSPM LE SLAT SW 28V DC STBY BUS 115V AC L MAIN BUS 28V DC L MAIN BUS

EADI

EFIS SYMBOL GENERATORS C

EICAS COMP

A/T DISC

a

TAT + 12C

AFCS WARN

+38C

COLLINS 96.1

A

12

GRD TEST SEL UP

TMC AC

POWER SUPPLY

DOWN

YES/ADV

ON/OFF

FLT FAULTS

GND TEST

8

26.1

4

8 N1

318

320 EGT

28V DC L MAIN BUS

TMC

MCDP

TMS - FUNCTIONAL DESCRIPTION - SYSTEM BLOCK DIAGRAM CONT. B767-3S2F Page - 135

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115.2 4

NO/SKIP

TMC SERVO

TMC P11 OVERHEAD CB PANEL

96.1

D-TO

54.9

EICAS DISPLAY UNIT

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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 TESTS

MODE CONTROL PANEL

FLIGHT/ GROUND FAULTS

LRU GROUND TESTS

SYSTEM GROUND TESTS

TEST 04 MCP

TEST 30 CURRENT FAULT REPORT TEST 40 AUTOLAND

AUTOLAND STATUS ANNUNCIATOR

TEST 06 ASA

SUPPORT GROUND TESTS TEST 51 AIR/GND RLY

AILERON SERVOS

TEST 07 SERVO AIL

ELEVATOR SERVOS

TEST 08 SERVO ELEV

TEST 56 FCC CONFIG/OPT FLIGHT CONTROL COMPUTER

TEST 59 FCC INSTR TEST 65 STAB TRIM

RUDDER SERVOS

TEST 09 SERVO RUDDER TEST 66 XDCR OUTPUTS TEST 67 AIL SURF LIM

AUTOPILOT DISCONNECT SWITCHES

TEST 11 SW A/P DISC

GO-AROUND SWITCHES

TEST 13 SW GA

TEST 68 ELEV SURF LIM TEST 69 RUD SURF LIM TEST 01 FCC

MAINTENANCE FUNCTIONS B767-3S2F Page - 137

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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 UP

YES/ADV

DOWN

NO/SKIP

MCDP GND TEST

(11 LINES)

ON/OFF

FLT FAULTS

TYPICAL CARRY-ON REMOTE MCDP CONTROL PANEL

MCDP P6-5 PANEL PLUG

EICAS COMPUTERS

LOWER EICAS DISPLAY UNIT

MAINTENANCE MONITOR SYSTEM - REMOTE MCDP OPERATION B767-3S2F Page - 139

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MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS Primary Interface Systems These are the primary interface systems: - Flight control computers - Flight management computers - Thrust management computers - EICAS computers - Remote MCDP control panel - Air/Ground relays . The left, center, and right flight control computers use ARINC 429 data buses 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.

Each flight control computer gets sensor data from its inertial reference unit and air data computer on ARINC 429 data buses. The thrust management computer gets control signals from the thrust mode select panel on an ARINC 429 data bus. The thrust management computer supplies analog control signals to the throttle servo motor and gets analog servo position signals. Ground Test Interface The primary systems are directly monitored by the MCDP. During the ground test, these systems interface with the secondary systems and various sensor/ control inputs.

AFCS MODE CONTROL PANEL

THROTTLE SERVO MOTOR GEN

ELEVATOR, AILERON AND RUDDER ACTUATORS (3) STAB TRIM DRIVES (2) AUTOLAND STATUS ANNUNCIATORS (2)

AIR/GROUND RELAYS (2)


AIR DATA/INERTIAL REF UNITS (3) AUTOPILOT DISCONNECT SWITCHES (2) GO-AROUND SWITCHES (2) SLAT SWITCHES FLAP TRANSDUCERS HYDRAULIC VALID OUT

FLIGHT CTRL COMPUTERS (3)

REMOTE MCDP CONTROL PANEL PLUG

FLIGHT MGMT COMPUTERS (2)

THRUST MGMT COMPUTER

MAINTENANCE CONTROL AND DISPLAY PANEL

MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS B767-3S2F Page - 141

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A/T DISC SW (2) PACK SWITCHES ANTI-ICE SWS AIR-DRIVEN PUMP SW ISLN VALVE SW REV THRUST SW SLAT SWITCHES FLAP TRANSDUCER

EICAS COMPUTERS (2)

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MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS CONT. Primary and Secondary Systems Faults in these LRUs can be reported by the MCDP: - FCC (3) - CDU (2) - FMC (2) - TMC - AIR/GND relays (2) - TMSP - MCP. System Interface The AFDS interfaces with these components: - Flap position - Stab position - ASA indicators - 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.

These are ARINC 700 sensors: - ILS - ADIRU - VOR - IRS - DME - Radar altimeter. Faults in these LRUs/systems are not reported: - Automatic direction finding (ADF) system - Communication radios - Air traffic control (ATC) system - Weather radar - Caution/warning system - Ground proximity warning system (GPWS) - M/ASI - Standby Altimeter - Yaw damper/YSM - Electronic flight instrument system (EFIS) - Vertical speed indicator (VSI) - RDMI.

AFCS MODE MODECONTROL CONTROLPANEL PANEL AFCS

THROTTLE THROTTLE SERVO SERVO MOTOR GEN MOTOR GEN

ELEVATOR,AILERON AILERONAND AND ELEVATOR, RUDDERACTUATORS ACTUATORS(3) (3) RUDDER STABTRIM TRIMDRIVES DRIVES(2) (2) STAB AUTOLANDSTATUS STATUS AUTOLAND ANNUNCIATORS (2) ANNUNCIATORS (2)

AIR/GROUND AIR/GROUND RELAYS(2) (2) RELAYS

THRUSTMODE MODE THRUST SELECT SELECTPNL PNL

AIR COMPUTERS AIR DATA DATA/INERTIAL REF (2) UNITS (3) AUTOPILOT DISCONNECT DISCONNECT AUTOPILOT SWITCHES (2) SWITCHES GO-AROUND SWITCHES GO-AROUND SWITCHES (2) (2) SLAT SWITCHES SLAT FLAP TRANSDUCERS TRANSDUCERS FLAP HYDRAULIC VALID HYDRAULIC VALID OUT OUT

FLIGHTCTRL CTRL FLIGHT COMPUTERS (3) COMPUTERS (3)

REMOTE MCDP REMOTE MCDP CONTROL CONTROL PANEL PLUG PANEL PLUG

FLIGHT FLIGHTMGMT MGMT COMPUTERS COMPUTERS(2) (2)

THRUST THRUSTMGMT MGMT COMPUTER COMPUTER

MAINTENANCE CONTROL AND DISPLAY PANEL MAINTENANCE CONTROL AND DISPLAY PANEL

MAINTENANCE MONITOR SYSTEM - INTERFACE SYSTEMS CONT. B767-3S2F Page - 143

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DISC A/TA/T DISC SWSW (2)(2) PACK SWITCHES PACK SWITCHES ANTI-ICE SWS ANTI-ICE SWS AIR-DRIVEN PUMP SWSW AIR-DRIVEN PUMP ISLN VALVE SWSW ISLN VALVE REV THRUST SWSW REV THRUST SLAT SWITCHES SLAT SWITCHES FLAP TRANSDUCER FLAP TRANSDUCER

EICAS EICAS COMPUTERS (2)(2) COMPUTERS

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MAINTENANCE CONTROL AND DISPLAY PANEL Purpose The purpose of the MCDP is to get and store fault data from the flight management system computers and to initiate, select, and process ground test functions.

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 commands the CPU to enter the ground test routine. GND TEST SEL UP Switch

Power Required

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 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 MCDP is a 6 MCU.

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, 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 mode.

Front Panel Features - Display Window The MCDP display has two lines of data with 16 characters on each line. Each character contains 17 segment gas discharge lamps, including decimal point. All displays are in the English language. Control Switches - ON/OFF 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 in the central processing unit (CPU). The power-up routine turns on the switch/ light. The switch/light goes off and the FAIL message is removed after five seconds if the MCDP passes the self-test. MODE SWITCHES 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.

NO/SKIP Operation Switch 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. Operating Instruction Plate The instruction panel is on the left side of the front panel. It contains instructions for these functions: - Apply power and initialize unit - Select mode of operation - Respond to interrogative messages - Ground test operation.

INFLIGHT FAULTS MODE SELECT SWITCH

MESSAGE DISPLAY WINDOW Collins

GROUND TEST NUMBER SELECT SWITCH (UP SLEW) GROUND TEST NUMBER SELECT SWITCH (DOWN SLEW)

GND TEST SEL UP

YES/ADV

POWER ON/OFF

DOWN

r

GRD TEST

NO/SKIP

w

GROUND TESTS LRU

2 SELECT MODE (FLT FAULTS OR GRD TESTS).

INSTRUCTION PLATE

w

INSTRUCTIONS 1 CYCLE POWER OFF/ON TO SELF TEST. VERIFY SEGMENTS LIGHT. VERIFY FAIL LIGHT OUT IN 5 SECONDS.

3 ANSWER QUESTION “?” WITH YES OR NO SWITCH. 4 FLT FAULT DATA: LAST FLIGHT FAULTS/PREVIOUS FLIGHT FAULTS. 5 GROUND TESTS AVAILABLE AS SHOWN: (NOTE: FAILURE EXISTS WHEN MESSAGE WILL NOT ADVANCE)

01 02 04 05 06 07 08 09

FCC TMC MCP TMSP ASA SERVO AIL SERVO ELEV SERVO RUD

SUPPORT 10 11 12 13

SERVO A/T SW A/P DISC SW A/T DISC SW G/A

14 15 16 17

XDCR COL L XDCR COL R XDCR WHL PVD/PVDC

SYSTEM 30 CURRENT FAULT REPORT 40 AUTOLAND

51 52 56 57 59 60 64 65 66 67 68 69

AIR/GRD RLY TMC RLY/SW FCC CONFIG/OPT TMC CONFIG/OPT FCC INSTR TMC INSTR SPD BRK/FLAP XDCR STAB TRIM XDCR OUTPUTS AIL SERVO LIMIT ELEV SERVO LIMIT RUD SERVO LIMIT

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YES/ADVANCE SWITCH

MODE FLT FAULTS

FAIL

POWER ON/OFF SELECT SWITCH

GROUND TEST MODE SELECT SWITCH


NO/SKIP SWITCH

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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 numbers and names for the LRU, SYSTEM, and SUPPORT tests that the operator can select.

- Auto on - Power up self-test - 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 flight management computers (FMC), and the thrust management computer on digital ARINC 429 receivers. It sends information to the same computers on ARINC 429 transmitters.

The flight faults are stored in an EPROM nonvolatile memory. The memory can be cleared in the shop with ITS test equipment.

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 FAULTS MODE SELECT SWITCH

MESSAGE DISPLAY WINDOW Collins

GROUND TEST NUMBER SELECT SWITCH (UP SLEW) GROUND TEST NUMBER SELECT SWITCH (DOWN SLEW)

GND TEST SEL UP

YES/ADV

POWER ON/OFF

DOWN

r

GRD TEST

NO/SKIP

3 ANSWER QUESTION “?” WITH YES OR NO SWITCH. 4 FLT FAULT DATA: LAST FLIGHT FAULTS/PREVIOUS FLIGHT FAULTS. 5 GROUND TESTS AVAILABLE AS SHOWN: (NOTE: FAILURE EXISTS WHEN MESSAGE WILL NOT ADVANCE)

GROUND TESTS

01 02 04 05 06 07 08 09

FCC TMC MCP TMSP ASA SERVO AIL SERVO ELEV SERVO RUD

SUPPORT 10 11 12 13

SERVO A/T SW A/P DISC SW A/T DISC SW G/A

14 15 16 17

XDCR COL L XDCR COL R XDCR WHL PVD/PVDC

SYSTEM 30 CURRENT FAULT REPORT 40 AUTOLAND

51 52 56 57 59 60 64 65 66 67 68 69

AIR/GRD RLY TMC RLY/SW FCC CONFIG/OPT TMC CONFIG/OPT FCC INSTR TMC INSTR SPD BRK/FLAP XDCR STAB TRIM XDCR OUTPUTS AIL SERVO LIMIT ELEV SERVO LIMIT RUD SERVO LIMIT

MAINTENANCE CONTROL AND DISPLAY PANEL CONT. TRAINING MANUAL ATA 22-00

B767-3S2F Page - 147

w

LRU

2 SELECT MODE (FLT FAULTS OR GRD TESTS).

INSTRUCTION PLATE

w

INSTRUCTIONS 1 CYCLE POWER OFF/ON TO SELF TEST. VERIFY SEGMENTS LIGHT. VERIFY FAIL LIGHT OUT IN 5 SECONDS.

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YES/ADVANCE SWITCH

MODE FLT FAULTS

FAIL

POWER ON/OFF SELECT SWITCH

GROUND TEST MODE SELECT SWITCH


NO/SKIP SWITCH

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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 B767-3S2F Page - 149

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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 B767-3S2F Page - 151

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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 IRU

FCC R

FAULT DATA

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

TMC

FAILED INTERFACE

MCDP FCC R (FMC L + FMC R + TMC) = IRU R FAILURE RIGHT IRU FAILURE FCC R

FAULT DATA

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 B767-3S2F Page - 153

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B767-3S2F Page - 154

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FAULT MESSAGE FORMATS FLIGHT FAULT MESSAGE FORMAT Historical Flight Tag and Fault Storage 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 and TMC. The STORE DATA Bit is transmitted by one of these computers only when the following sequence of events has occurred: IRS ground speed 100 KNTS and air/ground = in air, THEN IRS ground speed 40 KNTS and air ground = on ground ,Additionally, an engaged FCC channel will delay setting its STORE DATA bit until the channel is disengaged. 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.

Ground Test Diagnostics The diagnostic is displayed in the ground test mode (Test 30 CURRENT FAULT REPORT only) by pushing the GRD TEST switch when an LRU fault is being displayed. Content The diagnostic code defines the monitor within a computer that tripped to produce the displayed flight deck effect and faulty LRU. The decimal codes and their related monitors are defined in the following tables. An intermittent fault is defined as a fault that sometime in the past, during flight, caused a monitor to trip, but at some later time during the flight did not. No intermittent bit is present for ground test diagnostics.

FLIGHT FAULT MESSAGE FORMAT CHANNEL L=1 R=2 C=3

XXXXXX

INTERMITTENT OR SSFD 1 = YES (CASCADING) 0 = NO (VOTE OUT) EXAMPLE : DECIMAL CODE ASSIGNMENT (FROM TABLE)

SOURCE FCC = 1 FMC = 2 TMC = 3

INTERMITTENT OR SSFD

R 121541 FCC

DIAGNOSTIC CODE FORMAT HISTORICAL FLIGHT TAG 00 - PRESENT FLIGHT FAULT 01 - LAST FLIGHT FAULT 02-99 - PREVIOUS FLIGHT FAULT

HFT

FLIGHT DECK EFFECT OR DIAGNOSTIC

FDE OR DIAGNOSTIC FAULTY LRU FAULTY LRU OR TWO LRU(S) SEPARATED BY A / WHICH MAY MEAN AN INTERFACE FAULT

FAULT MESSAGE FORMATS B767-3S2F Page - 155

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