B767 - Flight Controls

B767/27/101 Flight controls

Boeing 767-200/300

Flight controls Training manual For training purposes only LEVEL 1

ATA 27

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B767/27/101 Flight controls

Training manual

This publication was created by Sabena technics training department, Brussels-Belgium, following ATA 104 specifications. The information in this publication is furnished for informational and training use only, and is subject to change without notice. Sabena technics training assumes no responsibility for any errors or inaccuracies that may appear in this publication. No part of this publication may be reproduced, stored in a retrieval system, or transmitted, in any form or by any means, electronic, mechanical, photocopying, recording, or otherwise, without the prior written permission of Sabena technics training.

Contact address for course registrations course schedule information Sabena technics training [email protected]

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TABLE OF CONTENTS 1. INTRODUCTION. ...................................................................................................................8 1.1. General................................................................................................................................8 2. AUTOFLIGHT INTERFACE. ..................................................................................................10 2.1. Hydraulic Shutoff Valves.....................................................................................................12 3. HYDRAULIC PWR SUPPLY. ..................................................................................................14 3.1. Hydraulic Distribution.........................................................................................................14 4. CSEU. ...................................................................................................................................16 4.1. General..............................................................................................................................16 5. AILERON..............................................................................................................................18 5.1. General..............................................................................................................................18 5.2. Aileron System Overview. ...................................................................................................20 5.3. Aileron Controls & Indications. ...........................................................................................24 5.4. System description. ............................................................................................................26 5.5. Aileron Trim Control...........................................................................................................30 6. SPOILERS. ............................................................................................................................36 6.1. General..............................................................................................................................36 6.2. System Operation. .............................................................................................................38 6.2.1. General. ...................................................................................................................38 6.2.2. Roll Spoiler. ..............................................................................................................40 6.2.3. Speedbrakes. ............................................................................................................42 6.3. Auto Speed Brake System. .................................................................................................44 6.3.1. General. ...................................................................................................................44 7. RUDDER. .............................................................................................................................46 7.1. General..............................................................................................................................46 7.2. Control & Indication...........................................................................................................48 7.3. Rudder System Description.................................................................................................50 7.3.1. General. ...................................................................................................................50 7.3.2. Rudder Hydraulic Distribution. ..................................................................................52 7.4. Rudder Trim. ......................................................................................................................54 7.5. Yaw Damper System. .........................................................................................................56 7.5.1. General. ...................................................................................................................56 8. ELEVATOR. ..........................................................................................................................58 8.1. General..............................................................................................................................58 8.1.1. Inboard and Outboard Elevator. ................................................................................60 8.2. System Description.............................................................................................................62 8.2.1. Elevator Schematic. ..................................................................................................62 8.2.2. Elevator Feel Computer. ............................................................................................64 8.2.3. Stick Nudger. ............................................................................................................66

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8.3. Position Indication..............................................................................................................68 8.3.1. Elevator System Travel...............................................................................................68 8.3.2. Elevator Position Transmitter. ....................................................................................68 9. STABILIZER. .........................................................................................................................70 9.1. General..............................................................................................................................70 9.2. Control & Indication...........................................................................................................72 9.3. System Description.............................................................................................................74 9.4. Stabilizer Travel Limits. .......................................................................................................76 9.5. Trim System. ......................................................................................................................78 9.5.1. Manual Electric Trim Switches. ..................................................................................78 9.5.2. Auto Trim. ................................................................................................................78 9.5.3. Mach Trim Mode. .....................................................................................................80 10. FLAPS & SLATS..................................................................................................................82 10.1. General............................................................................................................................82 10.2. System Operation. ...........................................................................................................85 10.3. Control & Indication.........................................................................................................88 10.4. Trailing edge FLAPS. .........................................................................................................90 10.4.1. General. .................................................................................................................90 10.4.2. Normal Operation...................................................................................................92 10.4.3. Flap Alternate Operation. .......................................................................................94 10.4.4. Flap Load Relief. .....................................................................................................96 10.4.5. Flap Disagree. .........................................................................................................96 10.4.6. Flap Asymmetry. .....................................................................................................97 10.5. Leading Edges Slats..........................................................................................................98 10.5.1. General. .................................................................................................................98 10.5.2. INB / OUTB Slats. ..................................................................................................100 10.5.3. Slat Position..........................................................................................................102 10.5.4. LE Slat Alternate. ..................................................................................................104 10.5.5. Slat Asymmetry and Failure Protection Shutdown. ................................................106 10.5.6. Slat Disagree. .......................................................................................................106 11. WARNINGS. ....................................................................................................................108 11.1. Stall Warning System .....................................................................................................108 11.1.1. General. ...............................................................................................................108 11.1.2. System Description. ..............................................................................................108 11.1.3. Components Description. .....................................................................................110 11.2. Takeoff Configuration Warning. .....................................................................................112 11.2.1. General. ...............................................................................................................112 11.2.2. Components Description. .....................................................................................114 11.3. Landing Configuration Warning. ....................................................................................116 11.3.1. General. ...............................................................................................................116 11.3.2. Components Description. .....................................................................................118 11.4. Speedbrake Warning. ..............................................................................................120

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LIST OF ILLUSTRATIONS AILERON - SYSTEM SCHEMATIC ........................................................................................ 21 AILERON CONTROL - SIMPLIFIED ........................................................................................ 27 AILERON CONTROLS & INDICATION ................................................................................... 25 AILERON DEFLECTION LIMITS ............................................................................................. 29 AILERON - LOCKOUT MECHANISM .................................................................................... 34 AILERON TRIM CONTROL ................................................................................................... 31 AUTO FLIGHT INTERFACE ................................................................................................... 11 AUTOSPEED BRAKE - GENERAL.......................................................................................... 45 CONTROL & INDICATOR .................................................................................................... 55 CSEU BLOCK DIAGRAM..................................................................................................... 17 LATERAL CONTROL ............................................................................................................ 19 ELEVATOR CONTROL – GENERAL ....................................................................................... 59 ELEVATOR CONTROL SCHEMATIC...................................................................................... 63 ELEVATOR FEEL COMPUTER ............................................................................................... 65 ELEVATOR POSITION .......................................................................................................... 69 FLIGHT COMPARTMENT CONTROLS & INDICATIONS.......................................................... 49 FLIGHT CONTROL ACTUATORS & SERVOS ........................................................................... 9 HIGH LIFT DEVICES ............................................................................................................ 83 HIGH LIFT SYSTEM ............................................................................................................. 87 HYDRAULIC DISTRIBUTION ................................................................................................ 15 HYD SHUTOFF VALVES ....................................................................................................... 13 HORIZONTAL STABILIZER .................................................................................................... 71 INBOARD AND OUTBOARD ELEVATORS ............................................................................. 61 LANDING CONFIGURATION WARNING............................................................................. 117 LANDING CONFIG. WARNING ......................................................................................... 119 LE SLAT ALTERNATE DRIVE CONTROL ............................................................................... 105 L.E. SLATS - GENERAL ........................................................................................................ 99 MANUAL & AUTO TRIM MODE .......................................................................................... 81 OUTBOARD SLAT DRIVE & TRACKS .................................................................................. 101 ROLL SPOILER DEFLECTION ................................................................................................ 41 RUDDER - GENERAL ........................................................................................................... 47 RUDDER HYDRAULIC DISTRIBUTION .................................................................................. 53 RUDDER - SYSTEM DESCRIPTION ....................................................................................... 51 SPEEDBRAKE DEFLECTION ................................................................................................. 43 SPOILERS - GENERAL.......................................................................................................... 37 SPOILER SYSTEM GENERAL ................................................................................................ 39 STABILIZER GENERAL ......................................................................................................... 75 STABILIZER TRAVEL LIMITS ................................................................................................. 77 STALL WARNING SYSTEM ................................................................................................ 109 STALL WARNING SYSTEM COMPONENTS ........................................................................ 111

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STICK NUDGER .................................................................................................................. 67 SYSTEM COMPONENTS ................................................................................................... 119 TAKE-OFF CONFIGURATION WARNING – WEU ................................................................. 113 TAKE-OFF CONFIG. WARNING SYSTEM COMPONENTS .................................................... 115 TE FLAP ALTERNATE DRIVE CONTROL................................................................................. 95 THS - CTL & INDICATION.................................................................................................... 73 TRAILING EDGE FLAPS ....................................................................................................... 91 WING COMPONENTS / DROOP MECHANISM ..................................................................... 33 YAW DAMPER SYSTEM ...................................................................................................... 57

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ABBREVIATIONS AND ACRONYMS A/P AC ADC AIDC ALTN AOA ARINC ATA BITE CG CMD COL CSEU CTL DC EFIS EHSV EICAS ELEC ELEV FCC FCT FLT FMC FSEU FSPM FSPM GND GPWC HYD IAS INB INOP IRS IRU ISLN KCAS LCCA LE LRU

Auto Pilot Alternate Current Air Data Computer Aircraft Integrated Data System Alternate Angle Of Attack Aeronautical Radio Incorporated Air Transport Association of America Build In Test Equipment Center of Gravity Command Column Control System Electronic Unit Control Direct Current Electonic Flight Instruments System Electro-Hydraulic Servovalve Engine Indicating and Crew Alerting System Electric, Electrical, Electricity Elevator Flight Control Computer Feel Centering & Trim Flight Flight Management Computer Flap/Slat Electronic Unit Flap Stabilizer Position Module Flap/Slat Position Module Ground Ground Proximity Warning Computer Hydraulic Indicated Airspeed Inboard Inoperative Inertial Reference System Inertial Reference Units Isolation aircraft speed (knots) Lateral Central Control Actuators Leading edge Line Replaceable Unit

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LVDT LVDT LVT Mc MCDP NVM OUTB OUTB(D) PCA PDU PPM PSEU psi PSM PTU PWR RRCM RVDT RVDT SAM SCM SPDBRK STCM SWC SYS TAS TE TMC UCM UNSCHD VAC VAL Vc WEU XDCR XMTR YDM

Linear Variable Differential Transducer Linear Variable Differential Transformer Linear Variable Transducer Mach number Maintenance Control & Display Panel Non Volatile Memory Outbound Outboard Power control Actuator Power Drive Unit Panel Position Monitor Circuit Proximity Switch Electronic Unit Pounds per Square Inch Power Supply Module Power Transfer Unit Power Rudder Ratio Changer Modules Runway Visual Range Rotary Variable Differential Transformers Stabilizer Trim & Aileron Lockout Modules Spoiler Control Modules Speedbrake Stabilizer Trim Control Modules Stall Warning Computer System True Airspeed Trailing Edge Thrust Management Computer Uncommanded Motion Unscheduled Voltage Alternating Current Valve Airspeed Warning Electronic Unit Transducer Transmitter Yaw Damper Modules

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1. INTRODUCTION. 1.1. General. Flight control systems can be grouped as primary or secondary control systems. Primary flight controls are those which are used to provide continuous control of the airplane about the pitch, roll and yaw axes, and include the aileron, rudder, elevator and spoiler systems. Secondary flight controls are those used intermittently, to modify the basic aerodynamic configuration of the airplane to improve its performance at a particular flight condition, and include the leading edge slat, trailing edge flap, spoiler (when used as air or ground speedbrakes) and stabilizer trim systems. Flight Control Actuators & Servos. All primary flight controls are driven by hydraulically operated Power Control Actuators (PCA) with no manual reversion capability. A total of twenty nine actuators are employed with eight in the aileron system, twelve in the spoiler system, six in the elevator system and three in the rudder system. In addition the aileron system has three additional control actuators (LCCA) to power the wing cable systems to the PCAs located at the aileron. Nine autopilot servos, three on each axis, provide triple redundancy required for category three autoland capability. The aileron servos are part of the LCCAs with the three elevator and three rudder servos as individual units. Two yaw damper servos provide rudder inputs independent of pilot/autoflight control inputs. The trailing edge flaps and leading edge slats which are secondary flight controls, are operated by power drive units (PDU). The PDU will rotate torque tubes to power two rotary actuators (mechanical) at each control surface.

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FLIGHT CONTROL ACTUATORS & SERVOS EFFECTIVITY ALL

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2. AUTOFLIGHT INTERFACE. Flight control computers (3) use the autoflight (A/P) servos to control airplane movement. FCC interfaces the thrust management computer (TMC), flight management computer (FMC), maintenance control and display panel (MCDP) with the specific A/P servo needed for airplane axis movement. FCC then commands the control valves in the appropriate servos to allow hydraulic pressure to move the output cranks. Output cranks connect to mechanical linkage for power control actuator (PCA) input. LVDT’s in each servo compare surface position vs servo position for the information needed to null commanded control valve inputs. Manual/electric overrides are available for each system.

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AUTO FLIGHT INTERFACE

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2.1. Hydraulic Shutoff Valves. Six Identical Shutoff Valves. Used to isolate flight control components from their hydraulic source during ground maintenance.

B767/27/101 Flight controls

Shutoff Valve Control & Indication. Six control switches located on the “HYD/GEN FIELD CONT” panel (P61). The switches are intended for ground use only and are normally on. These alternate action switches contain white “on” lights that are illuminated whenever the switch is in the open position.

Left and right wing shutoff valves are mounted on a bracket between the rear wing spar and the spoiler beam adjacent to the inbd corner of the inbd ailerons.

Each control switch is guarded by a cover which will not close in the switch off position (switch protruding).

Center wing shutoff valve is located on the aft bulkhead in the left wheel well.

The shutoff valve position is monitored by switch lights and EICAS messages.

Tail shutoff valves are located in the stabilizer compartment. Access is through a service door in the bottom of the compartment, just forward of the stabilizer jackscrew.

An amber light in the lower half of the control switch illuminates as soon as the valve moves from the fully open position.

The left and right tail shutoff valves are located approximately midway up on the second bulkhead ring forward of the access door.

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An amber EICAS advisory message appears on the upper display when a shutoff valve is not open (i. e. L WING HYD VAL). More than one valve not open will provide a single “FLT. CONT VALS” message.

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HYD SHUTOFF VALVES EFFECTIVITY ALL

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3. HYDRAULIC PWR SUPPLY. 3.1. Hydraulic Distribution. Three hydraulic systems operate power control actuators in the primary flight control systems which include the : - Aileron, - Elevator, - Rudder and - Spoiler systems.

Roll (LCCA), pitch (elevator) and yaw (rollout guidance) autopilot actuators are powered by each of the three hydraulic systems. Hydraulic shutoff valves control pressure to all flight control systems except the flaps, slats and stabilizer systems. The ram air turbine pump in the center hydraulic system can power all center hydraulic system flight controls except the flap and slat systems.

The secondary flight control-systems are powered by a combination of one, two or three hydraulic systems. - The stabilizer trim system is normally powered by the left and center hydraulic systems. A Power Transfer Unit (PTU) can provide right hydraulic system power to operate the left stabilizer trim system. - An elevator feel computer and yaw damper servo are powered by the left hydraulic system with identical components powered by the center hydraulic system. - The rudder ratio changer, the leading edge slats and trailing edge flap systems are powered by one hydraulic system.

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HYDRAULIC DISTRIBUTION EFFECTIVITY ALL

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4. CSEU. 4.1. General. Two identical control system electronic units (CSEU) are located in the main equipment center. Each CSEU contains eight modules which include two power supply modules (PSM) and six operating modules. The CSEU modules perform control, failure protection and fault indication functions for the : - Aileron, - Spoiler, - Stabilizer and, - Rudder systems. Modules are interchangeable between the left and right CSEUs. SCM. Spoiler control modules (SCM) - Each SCM receives command signals from rotary and linear variable differential transformers (RVDT/LVDT) and spoiler panel position from spoiler actuator LVDTs. Flap position is provided by three flap/stabilizer position modules (FSPM). Each SCM controls two actuators. One SCM in each CSEU receives a control wheel inhibit signal from the SAM.

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SAM. Stabilizer trim aileron lockout module (SAM) - The SAM receives stabilizer trim control inputs from flight control computers (FCC) and alternate and manual electric trim switches. Stabilizer and flap position inputs are from a FSPM. The SAM provide’s control signals to a stabilizer trim control module (STCM) . The SAM also uses speed signals from the air data system (ADS) and actuator position switch inputs to control two aileron lockout actuators. RRCM. Rudder ratio changer module (RRCM) - The RRCM receives speed signals from the SAM and position signals from a ratio changer mechanism LVDT to control the ratio changer actuator. YDM. Yaw damper module (YDM) - The YDM receives inputs from the air data and inertial reference systems and position signals from an actuator LVDT to control a yaw damper actuator. Hydraulic/Air Ground Inputs. All the CSEU operating modules receiver hydraulic pressure switch and air/ground relay signals for various control, test and fault indication functions.

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CSEU BLOCK DIAGRAM EFFECTIVITY ALL

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5. AILERON 5.1. General. The inboard and outboard ailerons provide roll control about the airplane longitudinal axis. The left and right ailerons move in opposite directions. The ailerons move up on one wing and down on the other, causing the airplane to roll. The inboard ailerons operate during all phases of airplane operation. The inboard ailerons are partially lowered with the trailing edge flaps to improve lift performance during takeoff and landing. The outboard ailerons are locked out during high speed flight to reduce roll sensitivity. The inboard ailerons are constructed of light weight composites. The skin is Graphite/Epoxy bonded to a core of Nomex Honeycomb. Ribs and spars are made of aluminum. Lower nose panels are removable for access to PCA’s and linkage. Attach points for the PCA’s are at the aileron midpoint. Four hinges attach the aileron to wing structure. The outboard hinges are fail-safe in design for improved safety margins. The outboard ailerons are constructed of light weight composite materials. The skin is made of Graphite/Epoxy bonded to a core of nomex honeycomb. Two adjustable tungsten balance weights prevent flutter if hydraulic power is lost. The aileron is attached to the wing structure by five hinges. Jumpers and static dischargers provide electrostatic protection to airplane systems.

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LATERAL CONTROL EFFECTIVITY ALL

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5.2. Aileron System Overview. The aileron control system provides mechanical control of hydraulic actuator servo valves. The system has two body cable systems to provide backup control for jams or disconnects. Lost motion devices provide separation between the primary and backup body cable systems. Three overrides are also installed between the body cable systems to provide separation of the systems if a jam occurs in either. Two overrides in each wing provide separation of a jammed inboard or outboard aileron from the aileron control system. Normal system operation by the primary cable system provides lateral central control actuator (LCCA) operation of the wing cable systems. Backup cable system operation of the LCCA’s is possible if a disconnect occurs in the primary cable system. If the LCCA control system is failed (primary system) direct operation of the right, wing cables by the backup cable system is possible by use of the override and lost motion devices. LCCA operation is required for left wing cable system operation. There are nineteen rig pins in the aileron system for adjustment of cables, rods, quadrants and torque tubes.

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RIGHT FWD CONTROL QUAD OVERRIDE 24 LBS LOST MOTION 6°

P61

TO R. WING

22LBS

STATIC

PITOT

PITOT

ADC FAIL EXT SENSOR FAIL

TEST

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B767/27/101 Flight controls

MECHANICAL STOP 65° CONT WHL SINGLE CHANNEL AUTOPILOT LIMIT IS 18° CONT WHL MULTI-CHANNEL AUTOPILOT LIMIT IS 55° CONT WHL AILERON TRIM LIMIT IS 30° CONT WHL

18LBS

STATIC

ADC FAIL EXT SENSOR FAIL

TEST

4° L. ADC

R. ADC

± 275 kts LOCK/UNLOCK

L. OUTPUT QUADRANT

L. LCCA’S

R. AIL. CTL QUADRANT - FEEL / CENTERING - TRIM MECHANISM R. LCCA’S LBS R. SAM

L. SAM

LOCK SIGN TO 3L SPOILER MODULE

14

TO R. OUTB. LOCKOUT ACT. LOCK SIGN TO 2R SPOILER MODULE

65°

4

T.E FLAPS DRIVE MECHANISM

4

65°

CONT WHL

M TRIM ACTUATOR

14

LOCKOUT ACTUATOR (2POS)

55° 65°

55° 65°

FLAP HANDLE DETENT

CONTROL WHEEL CW

5

25

30 UNITS

5 10 INB. AILERON DROOP POS

AILERON - SYSTEM SCHEMATIC

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5.3. Aileron Controls & Indications.

Control. Dual control wheels provide manual control of the aileron system. The wheels are mechanically connected by overrides and normally operate together. The ailerons are controlled during autoflight by switches on the mode control panel. Trim switches on the control stand, control an electric actuator which operates the aileron system. Indication. Aileron position is shown by pointers on the EICAS Status page. There is a trim indicator placard on top of each control column. Aileron lockout system faults are shown by an EICAS advisory message, an amber light and an EICAS maintenance message.

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AILERON CONTROLS & INDICATION EFFECTIVITY ALL

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5.4. System description. Forward Quadrants. Left and right forward control quadrants are operated by cables from the control wheels and are connected by a bus rod. The right quadrant contains override and lost motion devices. Left Wheel Well Area. Components located in or near the left wheel well include the feel, centering and trim mechanism and a torque tube connected to two lateral central control actuators (LCCA).

B767/27/101 Flight controls

General Operation. Control wheel inputs operate the primary cables which input to the feel, centering and trim mechanism. Linkages and torque tubes connect the feel, centering and trim mechanism to the LCCAs which operate the wing cables. The wing cables input into the aileron droop assembly which operates the PCAs for the inboard aileron. The wing cables also operate the outboard aileron lockout mechanism, which control the outboard aileron PCAs. A backup cable system is available for full or partial system operation if a disconnect or jam should occur in the primary cable system.

Right Wheel Well Area. A quadrant, operated by cables from the feel, centering and trim mechanism, connects to a torque tube which operates a single LCCA in the right wing. Two override devices are also located in this area. Wing Areas. Each wing contains an inboard aileron droop mechanism with two override devices, inboard aileron power control actuators (PCA), an outboard aileron lockout mechanism and outboard aileron PCA’s. Position transmitters are located near the outboard end of each aileron surface.

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AILERON CONTROL - SIMPLIFIED EFFECTIVITY ALL

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Aileron Deflection Limits. Aileron deflection is shown in relation to control wheel position. The outboard and inboard ailerons have different control wheel schedules. Maximum aileron deflection with no air load is determined by internal stops in the power control actuators except for the down displacement of the outboard ailerons. Outboard Ailerons. Maximum deflection is 30.5° up & 15.5° down at 50° of control wheel rotation. Down deflection is limited by the aileron lock mechanism. Inboard Ailerons. The inboard ailerons are lowered by the trailing edge flap system and thus have two control wheel schedules. Maximum deflection is 21.5 ° up and down, drooped or not drooped. When the inboard ailerons are not drooped, full up deflection requires 32° of control wheel rotation. Full up deflection requires 48° when the inboard ailerons are drooped. Autopilot. Maximum authority during single channel operation is 18° of control wheel rotation. Maximum authority is 55° when more than one autopilot is engaged during autoland operation. Aileron Trim. Aileron trim authority is 30° of control wheel rotation.

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AILERON DEFLECTION LIMITS EFFECTIVITY ALL

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5.5. Aileron Trim Control. The aileron trim arm and control switches are located on the aft end of the control stand. Both switches must be operated to power the trim actuator on the feel, centering and trim mechanism. Maximum aileron trim is 30° of control wheel rotation. Aileron trim indicator placards are located on top of each control column. Each unit of trim represents 5° of control wheel rotation.

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AILERON TRIM CONTROL EFFECTIVITY ALL

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Droop Mechanism. A droop mechanism is mounted inboard of each inboard aileron. This mechanism lowers or droops the inboard ailerons through trailing edge flap system operation of the aileron droop angle gearbox., The inboard and outboard droop quadrants have overrides for the inboard and outboard ailerons respectively. Both inboard ailerons are lowered (drooped) to improve lift when the trailing edge flaps are extended. The droop mechanism is mechanically operated by the aileron droop -angle gearbox as the flaps extend between up and 5 units causing the inboard ailerons to lower 10°. As the flaps are raised, the ailerons return to the neutral position. Extending the flaps from 25 to 30 units causes the inboard ailerons to retract approximately 5°.

B767/27/101 Flight controls

The outboard ailerons are locked out at high speed to reduce roll control sensitivity about the longitudinal axis of the airplane. An aileron lockout mechanism is mounted on the wing rear spar, inboard of each outboard aileron. The mechanism is operated by an electric actuator which positions linkages to prevent quadrant rotation from operating the control rod to the PCA’s. The lockout actuators are electronically controlled by stabilizer trim aileron lockout modules using speed inputs from the air data computers. Unlocked. Quadrant operation moves the idler lever which operates the drag link and the output crank to provide control rod output to the outboard aileron Power Control Actuator (PCA) control valves. Locked. Extension of the lockout actuator pulls on the actuator link which moves pivot point “All on the idler link over pivot point “B”. This moves the control rod and the outboard aileron to the neutral position. Quadrant operation causes the idler link to pivot only on point “A” without movement of the drag link or the control rod and the outboard aileron remains at the neutral position.

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WING COMPONENTS / DROOP MECHANISM EFFECTIVITY ALL

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AILERON - LOCKOUT MECHANISM EFFECTIVITY ALL

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THIS PAGE IS INTENTIONALLY LEFT BLANK

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6. SPOILERS. 6.1. General The twelve spoiler panels are operated as part of the lateral control system or as speedbrakes. When operated by aileron system input for lateral or roll control they are raised on one wing and lowered on the other. When operated by speedbrake mechanism inputs, to reduce lift and increase drag for descent and landing operations, they are raised on both wings.

Indication. Spoiler system faults cause and a maintenance message to be displayed. Faults causing auto shutdown of a panel pair cause display of the amber “SPOILERS” light and an advisory message. Addition fault information is available from the built in test function of the SCM.

There are two inboard and four outboard panels on each wing numbered from left to right. Spoiler operation is by electronic control of hydraulic actuators or a “fly-by-wire” control system. Control. Electronic control of hydraulic power actuators (PCA) is by control wheel rotary variable differential transformers (RVDT) and speedbrake lever linear variable differential transformer (LVDT) inputs to spoiler control modules (SCM). Each SCM outputs control signals to an electro-hydraulic servo valve (EHSV) on two power control actuators (PCA). Spoiler panel position signals from a piston operated internal PCA LVDT provide a feedback signal to the SCM for panel control and fault detection. RVDT lateral control inputs are from aileron system operation. LVDT speedbrake control inputs are from speedbrake lever operation. Linear variable differential transformers (LVDTs) within each PCA sends a feedback voltage back to the SCM. The command voltage and feedback are continually summed and the command voltage is nulled out as the spoiler approaches its commanded position. The panel stops when the feedback signal equals the command signal. EFFECTIVITY ALL

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SPOILERS - GENERAL EFFECTIVITY ALL

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6.2. System Operation. 6.2.1. General. Spoilers deploy as a function of control wheel deflection, speedbrake lever position, flap position, air/ground logic and aileron lockout logic. Each symmetrical pair of spoilers is controlled by a spoiler control module (SCM). Each spoiler is driven by a power control actuator (PCA). The output of each SCM and the feedback from the pair of PCAs being controlled by that SCM form a closed loop electrohydraulic servocontrol system.

B767/27/101 Flight controls

The actuator positions the LVDT pairs and speedbrake operation is accomplished through the SCMs as discussed above. Ground speedbrake deployment travel is determined by internal SCM programs. All three hydraulic systems are used for spoiler system operation. Flap position, air/ground and aileron lockout discretes are used for panel programming by the SCMs. Faults are annunciated on the EICAS displays and overhead P5 panel. Additional fault information is available using the SCM built in test function.

Control wheel rotation drives three rotary variable differential transformers (RVDT) in the left and right forward quadrants. Inflight speedbrakes are commanded by the speedbrake lever through three linear variable differential transformers (LVDT) pairs. Each SCM has two output signals - one to the left wing and one to the right wing. The two outputs will cause the spoiler pair to deploy asymmetrically during roll-control and symmetrically during speed brake operation. These modes are modified within the SCM if a roll is initiated while speedbrakes are deployed. Spoiler panel deployment is determined by internal SCM programs. Ground speedbrake operation may be initiated manually by moving the speedbrake lever to the deploy position or initiated automatically by an electrical actuator after the airplane is on the ground. Automatic speedbrake operation is controlled by switches and relays which command the auto speedbrake actuator when ARMING conditions are satisfied.

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SPOILER SYSTEM GENERAL EFFECTIVITY ALL

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6.2.2. Roll Spoiler. Operation of the lateral control system, by a control wheel or autopilot servo input, causes the spoilers to move up on one wing to roll the airplane. The amount of lateral control system input required to begin moving the panels up decreases when the flaps are at 25 or 30 units. A control wheel input, with the panels down, results in panels rising on the up aileron wing (down wing), and remaining down on the down aileron wing. A control wheel input with the panels full up, from a speedbrake lever input, results in panels lowering on the down aileron wing and remaining full up on the up aileron wing. A control wheel input with both wing panels partially raised results in panels rising on the up aileron wing and lowering on the down aileron wing. In this case the amount of panel movement would be reduced by one half to maintain the same rate of roll for a given control wheel input. Maximum deflection is 45° for the outboard panels and 17° for the inboard panels with full control wheel rotation.

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ROLL SPOILER DEFLECTION EFFECTIVITY ALL

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6.2.3. Speedbrakes. Speedbrake lever operation raises or lowers spoiler panels on both wings simultaneously. Panel deployment is proportional to the speedbrake lever position. The lever must be moved past the armed (8.5°) position to start raising the panels. The speedbrakes are full up when the lever is at 78°. During ground speedbrake operation, all panels deploy to a maximum of 60° For inflight speedbrake operation, outboard panels deploy to a maximum Of 45°. Inboard panels are limited to a maximum of 17°. Outboard panels 4 & 9 do not operate in the air from speedbrake lever inputs.

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SPEEDBRAKE DEFLECTION EFFECTIVITY ALL

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6.3. Auto Speed Brake System. 6.3.1. General. Manual Operation. A speedbrake lever on the center quadrant stand is connected to the LVDTs through a connecting rod and pivot shaft. The lever is lifted up and pulled aft to deploy speedbrakes. Speedbrake lever operation raises or lowers spoiler panels on both wings simultaneously. Panel deployment is proportional to the speedbrake lever position. The lever must be moved past the armed (8.5°) position to start raising the panels. The speedbrakes are full up when the lever is at 78°. During ground speedbrake operation, all panels deploy to a maximum of 60°. For inflight speedbrake operation, outboard panels deploy to a maximum of 45°. Inboard panels are limited to a maximum of 17°. Outboard panels 4 & 9 do not operate in the air from speedbrake lever inputs.

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Automatic Operation. Speedbrakes may be deployed automatically after the airplane lands. When all arming conditions are met, an auto speedbrake actuator extends and drives the pivot shaft through a no-back clutch. The clutch allows the pilot to override the actuator at any time. Automatic deployment requires the speedbrake lever in the armed position, thrust levers less than 8.5° from idle stop, and the airplane on the ground. If the lever is not in the armed position, movement of either reverse thrust lever into reverse will ARM the system provided the other arming conditions are satisfied. If a forward thrust lever is advanced beyond 8.5° from its idle stop, the auto speedbrake actuator will retract and return the spoilers to the down position. A switch cam monitors speedbrake lever position to activate the arming switch and the lever-position switch (takeoff configuration) . The cam also has a detent at the ARMED position for pilot feel. Reverse thrust linkage will lift the speedbrake lever out of the down detent and actuate the reverse thrust switch to arm the auto speedbrake actuator.

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AUTOSPEED BRAKE - GENERAL EFFECTIVITY ALL

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7. RUDDER. 7.1. General A single rudder on the aft spar of the vertical stabilizer provides yaw control about the vertical axis of the airplane. While the vertical stabilizer is the primary source of airplane directional stability, the rudder must also provide adequate directional control to coordinate turns, create sideslip, balance unsymmetrical engine thrust and enable landing during runway crosswind conditions. High engine thrust capabilities require a large rudder for directional control during engine failure on takeoff. Due to increased rudder effectiveness at high speed, rudder authority is reduced as speed increases to prevent structural damage. Movement of the rudder pedals is transferred by rods to forward quadrants connected to a pair of cables. The cables drive the aft quadrant. Autopilot servos and a rudder trim actuator also provide control inputs to the aft quadrant. Rudder control authority is varied from approximately 26° to 2° by the ratio changer mechanism. Rudder Ratio Changer Modules control the ratio changer actuator. A summing mechanism combines control inputs from the aft quadrant and the yaw damper servos. The servos are controlled by Yaw Damper Modules. The rudder is moved by three actuators, each powered by a separate hydraulic system. A position transmitter signals rudder position.

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RUDDER - GENERAL EFFECTIVITY ALL

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7.2. Control & Indication. Trim Control & Indication. A trim control knob on the control stand operates the electric trim actuator on the aft quadrant assembly. A trim indicator shows the trim actuator position in units of trim. Rudder Position indication. The rudder position is displayed on the EICAS status page. Yaw Damper Controls. The yaw damper system is controlled by two switches on the P5 overhead panel. An “ON” light shows the switch position. An amber “INOP” light indicates the yaw damper function is inoperative. A three position yaw damper test switch on the P61 panel tests both yaw damper systems. Warning Indications. A RUDDER RATIO amber light on the P5 overhead panel indicates the loss of rudder ratio changer function. EICAS caution and maintenance messages indicate various levels of ratio changer and yaw damper faults. Caution messages indicate loss of function. Maintenance messages indicate faults in the associated system.

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FLIGHT COMPARTMENT CONTROLS & INDICATIONS EFFECTIVITY ALL

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7.3. Rudder System Description. 7.3.1. General. Rudder Pedals. Pilot movement of the rudder pedals drives the aft quadrant assembly with a pair of cables. Maximum pedal movement drives the aft quadrant against the mechanical stop and moves the rudder 26.5° in each direction at low airspeed and no air load. Autopilot. Autopilot inputs cause the three directional rollout autopilot servos to drive the aft quadrant assembly. Autopilot inputs are provided in the autoland mode only and can command the rudder to approximately 23° Inputs backdrive the cable system to provide input to rudder pedal steering.

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Ratio Changer. The ratio changer mechanism provides a means of controlling the inputs to the PCAs by varying the input to the PCAs based on airspeed. The ratio changer actuator is controlled by rudder ratio changer modules. At the high speed position the ratio changer mechanism limits rudder movement to a maximum of 2.15°. Power Control Actuators (PCA). Three PCAs move the rudder each using a different hydraulic system. The left hydraulic system pressure to the middle PCA passes through the ratio changer actuator. If the ratio changer function is failed, the middle PCA is depressurized. Each PCA has an override in the input linkage to its control valve. Primary Control Path. A crush core load limiter prevents damage to the primary control path in the event of a system jam.

Trim. Trim inputs from the flight deck control switch drive the aft quadrant assembly and provide a maximum rudder movement of 16.8°. Trim operation backdrives the cables and rudder pedals.

Secondary Control Path. A spring override is provided in the yaw summing linkage to prevent damage to the secondary control path in the event of a system jam and to eliminate backlash of the dual path linkage.

Yaw Damper. Yaw damper inputs from two servos operate the yaw damper summing mechanism. These inputs are summed with other rudder control inputs. Yaw damper inputs do not backdrive the cable system. Each servo has a maximum authority of approximately 3° of rudder travel. There are shear rivets in the yaw damper servo output to the rudder control system.

A temperature compensating linkage, forward of the vertical stabilizer rear spar, functions to null control inputs from thermal expansion differences between control rods and fin structure.

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RUDDER - SYSTEM DESCRIPTION EFFECTIVITY ALL

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7.3.2. Rudder Hydraulic Distribution. Each autopilot servo and PCA is powered by a separate hydraulic system. The left hydraulic power to the middle PCA is routed through the rudder ratio changer actuator. The middle PCA is powered only when the ratio changer is operative. The ratio changer actuator is also powered by the left hydraulic system. A flow sensitive fuse (7 gpm) is installed in the left hydraulic system pressure line and a check valve (not shown) is in the left return line. The two yaw damper servos are powered by the left and center hydraulic systems. Hydraulic pressure to the rudder components can be shutoff by flight control shutoff valves controlled by switches on the P61 panel.

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RUDDER HYDRAULIC DISTRIBUTION EFFECTIVITY ALL

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7.4. Rudder Trim.

Trim Operation. Operation of the trim actuator backdrives the aft quadrant assembly and the rudder pedals.

Trim Control. The rudder trim knob located on the aft control stand controls power to the rudder trim actuator. Rotating the knob 5° removes a ground from the 15° & 25° switches. The knob must be rotated more than 25° to complete the circuit to the actuator.

When the landing gear is extended, hydraulic pressure is available to the nose wheel steering system. Rudder trim operation can then turn the nose wheels.

The actuator stroke is controlled by internal limit switches in the actuator motor. The actuator provides a trim authority of 67 percent of full rudder control resulting in 16.8° of rudder movement at low speed. The duty cycle of the trim motor is 30 seconds operation followed by 3 minutes cooling. Trim Indication. The rudder trim indicator is driven by a rotary variable differential transformer (RVDT) in the trim actuator. A pointer shows the rudder trim position in units. Full rudder trim should be indicated by a minimum of 14 units. An adjusting screw is on the aft side of the aileron/rudder control module to zero the pointer with the rudder centered and the feel, centering and trim mechanism rollers in the cam detent. Loss of power to the trim indicator causes the pointer to move off scale. The RVDT is not line replaceable.

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CONTROL & INDICATOR EFFECTIVITY ALL

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7.5. Yaw Damper System. 7.5.1. General. The yaw damper system provides automatic rudder control to improve airplane ride quality, dampen gust load on the vertical stabilizer, dampen undesirable sideslip and roll (Dutch roll) and coordinate turns. The yaw damper system includes two Yaw Damper Modules (YDM) each controlling a yaw damper servo actuator. The system uses inputs from the Air Damper Computers (ADC), the Inertial Reference Units (IRU), servo actuator LVTs, modal suppression accelerometers (installed on 767-300 only) and air/ground relays to command rudder movement. Pressure switches in the left and center hydraulic systems input for fault detection and indication.

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Controls & Indications. Two yaw damper control switches are located on the yaw damper control panel. The switches control engage power to the YDMs and to the yaw damper servo actuators. When a system is inoperative or during test, an INOP amber light illuminates in the switch and an advisory L (R) YAW DAMPER message appears on EICAS. A test switch is located on the P61 panel allows testing of the yaw damper system on the ground. Yaw Damper Modules. The YDMs command rudder movement and monitor yaw damper for faults. Failures are detected by automatic BITE and are stored in the module memory for ground recall. Faults are recalled with YDM face plate buttons and are indicated by displaying messages on a 12 character LED display.

Movements of the yaw damper servos are summed by a summing lever before transfer to the yaw summing mechanism which commands the rudder. When both yaw damper servos are operative in flight, maximum yaw damper input to the rudder is approximately 6° in each direction. When one servo is operative, maximum rudder movement is approximately 3°. The maximum amount of rudder command available for yaw damping depends on airspeed and the number of ADC and IRU supplying data.

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YAW DAMPER SYSTEM EFFECTIVITY ALL

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8. ELEVATOR. 8.1. General Two inboard and two outboard elevators are connected by hinges to the horizontal stabilizer rear spar. The inboard and outboard elevator on each side are connected together by connecting links and operate as a single unit. The outboard elevator is permanently rigged faired with the inboard elevator. Movement of the elevator provides primary control on the airplane pitch about the lateral axis. The purpose of the elevator is to make short term changes in the airplane pitch attitude for climb, descent and altitude hold. Elevator movement initiates long term trim by the horizontal stabilizer.

Fault Indication. The only elevator system fault indicated in the flight compartment is a fault of a feel computer output pressure. This fault is indicated by an ELEV FEEL message on the status and maintenance pages of EICAS. Elevator Position. Elevator positions are shown on the lower left corner of the EICAS status pages.

Moving the elevator trailing edge up results in an airplane nose-up movement. Moving the elevator trailing edge down results in an airplane nose-down movement. Control. Pilot control is by moving the control columns. Control cables transfer the input to elevator aft quadrants which command the PCA’s. The PCA’s are powered by the three hydraulic systems. In autopilot mode, the operating FCC electrically commands the autopilot servos which input to the elevator quadrants and the PCA’s. Control column feel is provided by the feel and centering unit. The feel at the column is varied as a function of airspeed by the elevator feel computer. A stick nudger commanded by the stall warning computers moves the elevators and the control columns to cause a nose down attitude of the airplane.

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ELEVATOR

LATERAL AXIS

PITCH CONTROL

ELEVATOR CONTROL – GENERAL EFFECTIVITY ALL

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8.1.1. Inboard and Outboard Elevator. Each elevator consists of an inboard and an outboard elevator fastened together by interconnecting links. The inboard elevator has three hinges and the attach point for the position transmitter. The outboard elevator has six hinges and three PCA attach fittings. The inboard and outboard elevators are removed separately. The inboard elevator is removed first and installed last. The inboard elevator weighs 97 lb and the outboard elevator weighs 199 lb. The elevators are built with graphite epoxy composite material.

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AFT VIEW

INBOARD AND OUTBOARD ELEVATORS EFFECTIVITY ALL

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8.2. System Description. 8.2.1. Elevator Schematic. The elevator control system is commanded by fore and aft movement of either control column, autopilot servos and stick nudger. The system has two redundant paths, one on the left side and one on the right side of the airplane. Each path commands its own elevator power control actuators (PCA). The two paths can be separated in case of a jam by overrides at the column torque tubes and at the aft quadrants.

B767/27/101 Flight controls

An override, mounted on the stabilizer compartment aft bulkhead, allows the elevator system to operate in case of a jam in the feel and centering unit. The three PCA’s on each side are commanded by the input control rod positioned by the aft quadrant torque tubes. They are powered by separate hydraulic system. Elevator movement is displayed on the EICAS status page responding to position transmitter input. The slave cable interconnect mechanism receives input from the elevator movement through a lost motion device and two overrides. The mechanism allows the elevator on one side to be commanded by the other in case of a lost connection downstream of the aft quadrant torque tube.

Movement of the column is transferred to a cable tension regulator quadrant. A stick shaker is mounted on each torque tube for stall warning. Cables connect the forward quadrants to the aft quadrants. The aft quadrants are mounted on their own offset torque tubes. These torque tubes are interconnected through an override and an asymmetry limiter. Three autopilot servos are connected to the torque tubes by connecting rods. A feel and centering unit, mounted on the stabilizer compartment aft bulkhead, provides feel and centering to the elevator system. The feel and centering is by spring and a dual hydraulic actuator. The feel computer, in the stabilizer compartment, varies the feel actuator hydraulic pressures as a function of airspeed. The stick nudger, mounted on the feel and centering unit, is commanded by the stall warning system to move the elevator down.

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CAPTAIN’S CONTROL COLUMN TENSION REGULATOR QUADRANT (2 PLACES)

F/O’S CONTROL COLUMN

OVERRIDE MECHANISM

FEEL AND CENTERING UNIT

AUTOPILOT PITCH CONTROL SERVO (3PL.)

CONTROL COLUMN OVERRIDE MECHANISM

SLAVE CABLE INTERCON.

HORIZONTAL STABILIZER STICK SHAKER (2 PLACES)

AFT QUADRANT INTERCON. ROD

ELEVATOR FEEL COMPUTER

AFT QUADRANT OVERRIDE MECHANISM

LEFT AFT QUADRANT OUTPUT ARM CONTROL ROD POSITION XMTR. (2PL.) CENTER LINE OF STABILIZER REAR SPAR HINGE

RIGHT AFT QUADRANT TORQUE TUBE

LEFT AFT QUADRANT TORQUE TUBE INTERCONNECT LINK

SLAVE CABLE QUADRANT (2 PLACES)

LEFT INBOARD ELEVATOR (RIGHT SIDE SIMILAR) PWR. CTL. ACTUATORS (PCA’s) (3 PL. ON EACH OUTBD ELEVATOR) LOST MOTION AND OVERRIDE DEVICE (2 PLACES)

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LEFT OUTBOARD ELEVATOR (RIGHT SIDE SIMILAR)

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8.2.2. Elevator Feel Computer. The elevator feel computer generates two variable hydraulic pressures based on pitot pressure and horizontal stabilizer position. The variable pressures provide feel forces at the elevator columns and operate the rate control valve of the Stabilizer Trim Control Modules (STCM).

B767/27/101 Flight controls

If a jam occurs in the right elevator controls, the captain feels one half of the hydraulic feel force, none of the mechanical feel force, and the override forces at the column and aft quadrant overrides. If a jam occurs in the left elevator controls, the first officer feels the same force as above plus the mechanical feel force.

The elevator feel computer is located on the left side of the stabilizer compartment. Operation. The elevator feel computer is a dual hydro-pneumatic unit which receives airspeed signal from two pitot pressure lines. The stabilizer position is transmitted to the feel computer by a rod connected to the stabilizer. As airspeed increases, the two output hydraulic pressures gradually increase. The horizontal stabilizer movement to trim the airplane nose-up gradually limits the output pressures to a lower value. The output feel pressures can vary from 175 psi to 1150 psi above return line pressure at airspeeds from 0 to about 350 knots with the horizontal stabilizer in the neutral position. A relief valve is provided in each half of the feel computer to limit feel pressures to 160% of normal output pressures. The output pressure from each computer half is monitored by two differential pressure switches. When one output feel pressure differs from the other by 25% or more for more than 30 seconds with all three hydraulic system pressurized, the ELEV FEEL message appears on the status and maintenance pages of EICAS. This message is latched in the air mode in the EICAS memory. A bias spring biases the stabilizer input crank towards the airplane nose-down position. If the stabilizer input rod fails, the feel pressures can then vary to the full range of pitot pressure changes.

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ELEVATOR FEEL COMPUTER EFFECTIVITY ALL

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8.2.3. Stick Nudger. The stick nudger consists of an electric actuator, crank assembly and spring mounted on the upper face of the feel and centering unit. The electric actuator is pivot mounted on the feel unit and rotates a crank assembly which then pulls on the stick nudger spring. The spring is attached to the upper (captain) input crank of the feel unit. The stick nudger actuator consists of a 28V DC electric motor which extends and retracts a rod. Switches internal to the actuator cutout electric power at the end of the stroke. Operation. During normal operation, the nudger actuator is retracted. The center line of the crank assembly left end coincides with the pivot of the upper input crank. Elevator command inputs are not affected by the stick nudger. During flight with flaps and slats retracted, a stall warning from both stall warning computers results in commanding the actuator to extend. As the actuator extends, the rotation of the crank assembly pulls on the nudger spring and rotates the upper input crank. If the pilot does not restrain the movement of the column, the stick nudger applies an airplane nose-down command to the elevator. If the pilot restrains the movement of the column, the stick nudger spring is stretched further as the nudger actuator extends. When the stall warning stops, the nudger actuator is commanded to retract and repositions the crank assembly.

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STICK NUDGER EFFECTIVITY ALL

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8.3. Position Indication. 8.3.1. Elevator System Travel. Maximum travel of elevator is limited by internal stops in the PCA’s. The control column stops and aft elevator quadrant stops provide backup elevator travel limitation after cable stretch. With no air load, the elevator can move a maximum of 28.5° up and 20.5° down with a full forward and aft movement of the control column. The elevator displacement is measured by the distance between the inboard elevator trailing edge and the reference point on the index plate riveted on the airplane tail cone with the stabilizer set at 2 units. At maximum travel of the elevator, these measurements are 33.7 in up and 24.5 in down. The maximum movement of the elevator in autopilot mode depends upon the number of flight control computers engaged. In single autopilot, the maximum authority is 8.3° up and down. In multiple autopilot, the maximum authority is 28° up and 20° down.

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8.3.2. Elevator Position Transmitter. The elevator position transmitter is mounted on the inboard side of the horizontal stabilizer on both sides of the airplane. The transmitter is powered by 28V AC and supplies a signal to the EICAS computers to display the elevator position pointers on the status page. Transmitter Failure. If a transmitter is failed, its position pointer is not shown on the status page. The elevator operation must be visually checked for each dispatch of the airplane. Full elevator movement on the ground should move the pointer to or past the up or down mark on the status page. Transmitter Replacement. The replacement procedure of the elevator position transmitters is the same as the ailerons transmitters. The transmitter is installed with the stabilizer set at 2 units and the elevator faired. The transmitter is adjusted to obtain a signal of less than 50 millivolts by the null adjusting sleeve on the control rod. The transmitter can also be adjusted by centering the pointer on the status page.

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ELEVATOR POSITION EFFECTIVITY ALL

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9. STABILIZER. 9.1. General Purpose. The horizontal stabilizer is a moveable assembly that includes the elevator. Changes in the stabilizer angle of attack result in airplane movement about the pitch axis. The purpose of the stabilizer trim is to make long term changes in the airplane pitch attitude (short term pitch changes are made by the elevator). Airplane pitch requirements change during flight due to changes in center of gravity (CG), engine thrust and airspeed changes. Moving the stabilizer leading edge up results in airplane nose down trim. Moving the stabilizer leading edge down results in airplane nose up trim.

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HORIZONTAL STABILIZER EFFECTIVITY ALL

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9.2. Control & Indication. Fault Messages & Lights. When the stabilizer moves without command or moves in the opposite direction than the one commanded, the caution (B) level message UNSCHED STAB TRIM appears on EICAS and the amber light UNSCHED STAB TRIM illuminates on the P5 panel. When only half rate trim results from, a pilot manual electric trim input, the advisory (C) level message STAB TRIM appears on EICAS and the amber light STAB TRIM illuminates on the P5 panel.

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9.3. System Description. General Description. Pitch trim of the airplane occurs when the stabilizer is moved leading edge up or down. Hydraulic power drives a ballscrew, attached to the stabilizer center section front spar, which causes it to rotate about aft hinges that are connected to fuselage structure.

B767/27/101 Flight controls

SAM. These modules get signals from the pilot, flight control computer (FCC) and air data computer (ADC). Using these inputs they control all modes of stabilizer operation except for alternate electric trim. In this mode signals are sent directly to the STCMs.

Power is supplied from the left and center hydraulic systems to two stabilizer trim control modules (STCM). Flow from these modules to two hydraulic motors is controlled by electrical inputs. Airspeed changes sensed by a elevator feel computer modifies trim rate.

Limit Switch & Position Transmitter Modules (3). The modules are designated left, right and center and are driven by cables when the stabilizer moves. They contain limit switches and position transmitters that set maximum stabilizer travel and provide the indication of stabilizer movement. Outputs from the modules are sent to three flap/stabilizer position modules (FSPM) that return signals to SAM and the FCC.

Manual Electric Trim. When either pilot moves the arm and control trim switches on the control wheel, inputs go through both SAMs to both STCMs.

Electric Alternate Trim Control. Two electric alternate trim switches, on the control stand, directly power the CONTROL & ARM solenoid valves of both STCMs.

The arm signal is directed through cutoff switches, controlled by elevator column movement, and limit switches. These switches are in the left/right position transmitter modules that are cable driven by stabilizer movement. Trim rate is maximum in this mode. Autoflight Trim. FCC auto stabilizer trim results from elevator inputs to the selected FCC. When this occurs auto trim signals are sent to one SAM. The control and arm outputs that result use the same path as described under manual electric trim. If the pilot or FCC is not making a trim input, mach trim circuits in the SAM will control trim operation using AIDC inputs. Trim rate in these modes is half a maximum as only one STCM is controlling one hydraulic motor/brake combination. EFFECTIVITY ALL

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LIMIT SWITCHES & POS. TRANSMITTER MODULES (3) MAN. TRIM

L

FCC C R

MAN. TRIM AUTO TRIM

AUTO TRIM ADC

L. HYD. SYST. & PTU PRESS. RET.

MACH TRIM

MACH TRIM

C. HYD. SYST. PRESS. RET.

R. SAM

L. SAM CUTOFF

TRAVEL LIMIT SWITCHES

ELEVATOR FEEL PRESS.

S.O.V.

CONTROL MODULES

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9.4. Stabilizer Travel Limits. The full movement of the stabilizer is from 0 unit to 14.2 units, one unit is equal to approximately 1°. The neutral position of the stabilizer is at 2.0 units. At the neutral position the stabilizer is aligned with the longitudinal axis of the airplane.

B767/27/101 Flight controls

Stabilizer Position Dimensions. The position of the horizontal stabilizer is checked by measuring a dimension A between the upper stop and the ball nut. The table shows dimensions relative to three basic referenence positions, electric limit and green band measurements are shown in the maintenance manual. The dimensions will vary depending on model and engine type.

Travel Limits. Travel limiting devices control the range of stabilizer movement in all operating modes. Upper and lower mechanical stops limit the stabilizer travel at each end of the range. In the electrical control mode, limit switches prevent stabilizer travel beyond that required by the normal flight envelope. The upper limit in the SAM electrical mode is provided at 0.25 units with the flaps extended and 1.5 units with the flaps retracted. The lower limit is provided at 12.8 units. Stabilizer trim limits for takeoff are between 0.25 units and 7.0 units. These limits are indicated by a green band on the pilots stabilizer position indicators. Three painted marks on the fuselage indicate stabilizer position.

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9.5. Trim System. 9.5.1. Manual Electric Trim Switches. A set of dual switches, located in each pilots wheels, controls stabilizer trim in the manual electric mode. Each switch set provides manual electric ARM and CONTROL trim signals to both STCMs through both SAMS. The switches can be moved up or down and are spring loaded to the OFF neutral position. Operation. When the switches on either wheel are moved together in an up or down direction, 28V DC power is directed through each SAM to the appropriate solenoids on each STCM. The auto manual transfer relays in each SAM are released during manual trim. When one FCC is engaged, the relays in both SAMS are energized allowing the auto trim commands of the SAM in of control to power the solenoid valves its dedicated STCM. If a manual electric trim command is made, coincidence monitors in the SAMS, inhibit the auto trim mode, disengage the FCC, and de-energize, both automanual relays. Now the manual electric trim command takes priority over auto trim. When there is multi channel FCC lays in engagement, the auto transfer relay in both SAMS are energized allowing auto trim commands from the SAM in control to power the solenoid valves of its dedicated STCM. If a manual electric trim command is then made, the coincidence monitor signals inhibiting the auto trim mode are ignored by the FCCs. The relays remain energized disabling the control of the manual electric trim switches.

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9.5.2. Auto Trim. Auto trim control of the stabilizer occurs when one or more FCC is engaged. The left FCC commands the left SAM and the right FCC commands the right SAM. The center FCC commands either the left or the right SAM depending on which SAM is initialized first. Operation. A selected FCC sends an ARM signal to its corresponding SAM. When it receives a VALID signal from the SAM, the FCC engages. The engaged FCC provides a trim command to the SAM when the elevator is deflected out of its neutral position. Single Channel FCC. During single channel operation, one FCC is engaged and commands its corresponding SAM. The auto-manual transfer relays are energized in both SAMS. The FCC provides an auto trim command when the elevator is deflected for more 4 seconds. The stabilizer then moves at half rate speed to new trim position. Multiple Channel FCC. During multiple channel operation, two or three FCCs are engaged. The FCC (first engaged) provides trim commands to its SAM, immediately, when the elevator is deflected out of its neutral position. The stabilizer moves at half rate speed. If a failure occurs on the 767-300 that results in autoland 3 (three FCCs engaged) defaulting to autoland 2 (two FCCs engaged), the remaining FCC not controlling trim engages the SAM not previously performing the auto trim function. The two FCCs command the two SAMS to trim simultaneously resulting in the stabilizer moving at full rate. This full speed trim permits quick retrimming of the airplane for go-around during a missed approach.

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9.5.3. Mach Trim Mode. General Description. The Mach trim mode controls the stabilizer trim during flight when no autopilot is selected (no FCC engaged) and no other stabilizer trim command exists. The controlling SAM commands the stabilizer trim as a function of change in Mach number to enhance the longitudinal stability of the airplane. A speed increase results in an airplane nose up trim while a speed decrease results in an airplane nose down trim. Operation. Both SAMs contain the Mach trim circuits and receive Mach data from the two ADCs. After initialization the controlling SAM automatically engages the Mach trim mode and commands the stabilizer to move at half rate of speed when : - The airplane is in the air (20 sec. delay), - No manual electric trim switches input is being made, - No alternate electric trim switches input is being made, - No FCC is engaged (no autotrim), - Flaps and slats are retracted, - The airplane speed changes, - No SAM fault is present. The Mach trim schedule provides for a greater stabilizer trim correction as the Mach number increases. Movement of the elevator control column in opposite direction to the airplane longitudinal trim causes the elevator control column cutoff switches to stop the Mach trim mode.

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MANUAL & AUTO TRIM MODE EFFECTIVITY ALL

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10. FLAPS & SLATS. 10.1. General High Lift Devices. The high lift devices include the trailing edge flap and leading edge slat systems. The high lift devices are extended to improve wing lift and drag characteristics for takeoff and landing operations and to provide increased stall operating margins.

The leading edge slats have three positions of retracted (up), intermediate (takeoff /sealed) , extended (landing/ gapped) and operate between lever, or switch, positions of up to one and 20 to 25. The alternate position selector switch has a NORM (normal) position which does not provide an output command. This is a safety position to prevent inadvertent alternate drive operation in event either arm switch is actuated and flap or slat position disagrees with the selector switch.

There are four trailing edge flaps which have six operating positions. The inboard flaps have main and aft sections and are double slotted when fully extended. The outboard flaps have one section and are always single slotted when extended. There are five outboard and one inboard slat surfaces on each wing. The slats are numbered from the left to the right wingtip. The slats extend from the top of the wing and have three positions. A Krueger seal flap extends from the bottom of each wing between the inboard slat and the engine strut to reduce drag with the slats extended. The flaps and slats are controlled by the flap control lever during primary (hydraulic motor) operation and by the alternate flap selector and arm switches during alternate (electric motor) operation. Alternate operation is about six times slower than normal operation. The flap lever and alternate position selector switch have positions showing units of trailing edge flap extension. The trailing edge flaps are retracted at both the up and one positions of the lever or switch.

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HIGH LIFT DEVICES EFFECTIVITY ALL

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High Lift Systems. The leading edge slat and trailing edge flap systems are usually operated together by the flap control lever. The trailing edge flaps are operated by one drive system and power drive unit (PDU) with two rotary actuators at each flap. The leading edge slat system has separate drives for the inboard and outboard devices with the two ‘inboard slats operated by one drive and PDU, and the ten outboard slats operated by another drive and PDU. There are two rotary actuators at each slat. A flap/slat electronic unit (FSEU) provides position indication, failure protection and control functions. A flap/slat shutoff valve module is controlled by the FSEU to sequence flap and slat operation. The FSEU also provides separate flap and slat alternate operation for nonnormal high lift systems operation. FSEU. The flap/slat electronic unit (FSEU) controls the slat shutoff valve during hydraulic motor operation to sequence the flaps and slats and for long-term system depressurization. The FSEU controls the PDU bypass valves for hydraulic motor shutdown for uncommanded motion and asymmetry failures. PSEU. The proximity switch electronic unit (PSEU) monitors proximity sensors on each slat to detect disagree and asymmetry faults. The PSEU receives, flap lever and alternate position selector switch inputs from the FSEU for disagree fault detection. The PSEU signals the FSEU when either fault is detected.

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10.2. System Operation. TE Flap Hydraulic Operation. Flap control lever, or load relief actuator, operation of the power drive unit (PDU) control unit input cam, moves the control valve module control valve from the null position. Center hydraulic system pressure is then provided to the hydraulic motor through the control valve module bypass valve. Flap drive shaft rotation operates the control unit follow-up cam to return the control valve to null. When the control valve is at null and the bypass valve is at normal, there is a hydraulic lock on the motor to hold the gearbox and flap drive. Hydraulic pressure to the motor can be shutoff by the Flap/Slat Shutoff Valve module flap shutoff valve or by the bypass valve in the PDU control valve module. If pressure is shutoff at the Flap/Slat Shutoff Valve Module the hydraulic lock remains on the motor. If pressure is shutoff by the bypass valve the motor can be rotated by the gearbox. Pressure is shutoff to the motor as follows : - Flap/slat shutoff valve module : The flap solenoid valve is powered by the flap/slat electronic unit (FSEU) to close the flap shutoff valve when the flaps and flap lever are up or, on extension, until the leading edge slats have moved to their intermediate position. - Bypass valve : The bypass valve is positioned to bypass when the alternate flap system arm switch is actuated or by the FSEU during flap system failure.

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LE Slat Hydraulic Operation. Flap lever operation of the power drive unit (PDU) pilot input arm operates the control unit input cam, to move the control valve module control valve from the null position. Center hydraulic system pressure is then provided to the hydraulic motor through the control valve module bypass valve. Slat drive shaft rotation operates the control unit follow-up cam to return the control valve to null. When the control valve is at null and the bypass valve is at normal, there is a hydraulic lock on the motor to hold the gearbox and slat drive. Hydraulic pressure to the motor can be, shutoff by the flap/slat shutoff valve module slat shutoff valve or by the bypass valve in the PDU control valve module. If pressure is shutoff at the flap/slat shutoff valve module, the hydraulic lock remains on the motor. If pressure is shutoff by the bypass valve, the motor can be rotated by the gearbox. Pressure is shutoff to the motor as follows : - Flap/slat shutoff valve module : The slat solenoid valve is powered by the flap/slat electronic unit (FSEU) to close the slat shutoff valve when the slat drive and flap lever are in agreement or, on retraction, until the trailing edge flaps are up. - Bypass valve: The bypass valve is positioned to bypass when the alternate slat system arm switch is armed or by the FSEU during slat system failure.

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Flap System Electronic Interface. Hydraulic Motor Operation. Failure protection shutdown, flap/slat sequencing, long-term pressure shutoff, and load relief during hydraulic motor operation are provided by flap slat electronic unit (FSEU) section I through control of the flap shutoff valve, load relief actuator and bypass valve. These control functions require inputs from a flap lever rotary variable differential transformer (RVDT), a flap position transmitter RVDT (from a Flap stabilizer position module (FSPM)), flap position transmitter resolvers and slat power drive unit (PDU) RVDTs. Alternate arm switch input inhibits failure protection shutdown and load relief functions in FSEU 1. Electric Motor Operation. FSEU section 3 controls the electric motor using inputs from the alternate flap selector switch and a flap position transmitter RVDT (from a FSPM). The flap alternate arm switch operates the PDU bypass valve and engages the electric motor clutch. Position Indication. Flap position transmitter syncros operate the flap indicator syncros. PSEU section 3 controls a flap reference transfer relay using inputs from a flap position transmitter RVDT (from a FSPM) and slat PDU RVDTs. Fault Annunciation. A fault light and messages are control by FSEU section 2 using inputs from a flap lever RVDT, alternate arm and position selector switches, flap position transmitter RVDT (from a FSPM) and slat position (from FSEU sections 1 and 3). FSEU section 1 controls a single fault message and inputs to section 2 for illumination of the associated fault light.

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HIGH LIFT SYSTEM EFFECTIVITY ALL

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10.3. Control & Indication. TE Flap Control & Indication. The flap control lever provides control during primary (hydraulic motor) operation of the flap system. Alternate (electric motor) operation is controlled by the ALTN FLAPS position selector and arm switches. Two needles on the flap position indicator show left and right wing flap drive positions. Flap drive position is shown in units with indicated airspeed (IAS) limits shown for each extended position.

B767/27/101 Flight controls

LE Slat Control & Indication. The flap lever provides control during primary (hydraulic motor) operation of the slat system. Alternate (electric motor) operation is controlled by the ALTN FLAPS position selector and arm switches. Two needles on the flap position indicator show slat position at the up and one unit positions. A LEADING EDGE amber light, master CAUTION lights and EICAS caution, and status and maintenance messages show slat system faults.

A TRAILING EDGE amber light, master CAUTION lights and EICAS caution alert, advisory alert, status and maintenance messages show flap system faults.

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10.4. Trailing edge FLAPS. 10.4.1. General. Two trailing edge flaps are mounted on each wing. The inboard flaps are double slotted and the outboard are single slotted. A single power drive unit (PDU) powers eight rotary drive actuators (two on each flap assembly) through gearboxes and drive shafts. Control of primary (hydraulic) flap operation is by a flap control lever connected by cables to the flap aft quadrant. A load relief (alleviation) actuator is connected to the PDU input linkage to limit extension if airspeed limits are exceeded at landing flap settings. A flap slat electronic unit (FSEU) controls position indicating, failure protection and alternate (electric) operations. Position transmitters on the flap drive actuators and flap control lever are used by the FSEU for its control functions. The FSEU controls the flap/slat shutoff valve module to sequence flap and slat drive operation and to remove hydraulic pressure to the PDU during cruise flight operations.

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TRAILING EDGE FLAPS

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10.4.2. Normal Operation. Flap Control Lever. The spring loaded flap lever has seven detented positions. The detents show units of primary drive command input. Gates at the land 20 unit positions prevent lever movement directly through these positions. These gates show lever positions for critical flap and slat configuration changes during flight operations. Two rotary variable differential transformers (RVDT) are operated by a flap lever gearbox in the control stand. RVDT number 1 inputs flap lever position to flap slat electronic unit (FSEU) section 2 and RVDT number 2 to inputs FSEU section 1. Flap Primary Drive Control. The hydraulic motor on the flap power drive unit (PDU) is the primary power source for the flap drive. Inputs to the PDU control unit for hydraulic motor operation are from cables and control rods operated by the flap control lever or by operation of the control rods by the flap load relief actuator. Drive shaft operation of the control unit provides hydraulic motor shutdown at the commanded position (closed loop). Cables from the flap control lever are routed through the forward cargo compartment to a flap aft quadrant in the aft cargo compartment. A control rod from the quadrant extends through the right wheel well aft wall. Controls shafts connect the control rod to the PDU. Turnbuckles are provided at two locations for cable rigging.

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10.4.3. Flap Alternate Operation. The flap power drive unit (PDU) alternate drive electric motor is controlled by the flap slat electronic unit (FSEU). The FSEU receives command inputs from the alternate flaps position selector switch. Alternate flap arm switch inputs are also used by the FSEU for relay control as well as for direct control of the PDU bypass valve. The FSEU controls alternate flap relays to provide power to the reversible PDU electric motor. A right wing drive position transmitter inputs to a flap stabilizer position module (FSPM) which provides flap drive position to the FSEU. This provides an FSEU closed loop control system, similar to primary drive, to turn off the electric motor when flap drive and selector switch inputs agree.

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TE FLAP ALTERNATE DRIVE CONTROL EFFECTIVITY ALL

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10.4.4. Flap Load Relief. The load relief (alleviation) system prevents excessive airloads on the flaps by automatically limiting flap extension when airspeed is too high for flaps 25 or 30. The system will limit flap position to a maximum of 20 units when the airspeed equals or exceeds 172.5 knots with the flap lever at 30 units or 182.5 knots with-the flap lever at 25 units. Operation requires a valid airspeed input from either air data computer (ADC). Load relief system operation is inhibited when the alternate flap system is armed or for center flap stabilizer position module (FSPM) failure, flap lever transmitter (RVDT) failure or loss of 28 volt ac power supply. Load relief operation is latched after actuation and is reset by reducing airspeed to 168.5 knots for flap extension to 30 units and to 178.5 knots for extension to 25 units. Airspeed must be valid from either ADC to reset. If both ADC inputs are invalid the latches can be reset by moving the flap lever to 25 or 20 units. A trailing edge amber light and a FLAP LOAD RELIEF advisory message will illuminate if the system fails when load relief is required. Disagreement annunciation between flap lever and flap position and the flap failure protection shutdown system are inhibited during load relief operation. Load relief operation is controlled by Section 1 of the flap slat electronic unit (PSEU), fault annunciation is controlled by Section 2.

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10.4.5. Flap Disagree. The FSEU provides a TE FLAP DISAGREE caution alert message and TRAILING EDGE amber light when the flap drive is not in the commanded position and the applicable time delay has expired. Different time delays provide for different operating times between flap positions and between the hydraulic and electric motors. Normal Operation. During normal flap system operation the hydraulic motor is controlled using the flap lever. The flap slat electronic unit (FSEU) Section 2 compares flap lever position to flap drive position from a flap stabilizer position module (FSPM). A disagreement is inhibited if the slats are less than takeoff, load relief is operating, a slat or flap power drive unit (PDU) is moving towards the lever position, an asymmetry condition exists or the alternate drive system is armed. Alternate Operation. During alternate flap system operation the electric motor is controlled using the alternate flap position selector switch. When the alternate flap arm switch is armed the FSEU compares the position selector switch to flap drive position from the FSPM. If the arm switch is armed for seven seconds with the position selector switch in NORM the message and light are illuminated to indicate both normal and alternate systems are disabled. The TRAILING EDGE amber light is also illuminated if a flap asymmetry condition is detected or the flap load relief system fails to operate when activated.

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10.4.6. Flap Asymmetry. The flap asymmetry protection system shuts down (bypasses) the flap primary drive and illuminates the TE FLAP ASYM message and flap amber light when a flap drive asymmetry condition is detected. Resolvers in flap position transmitter assemblies one and eight and four and five are compared by the flap slat electronic unit (FSEU) section 1 to detect a disconnect in the flap drive system. A difference in resolver degree input between resolvers one and eight or between four and five equivalent to 43 1/4 drive shaft revolutions causes immediate system shutdown and fault annunciation (approximately 12% of flap drive full travel). If the resolver difference remains for five seconds the fault and annunciation are latched. If the asymmetry no longer exists, the latch can be reset by pushing the flap alternate arm switch on and off, by opening and closing the FSEU 1 control circuit breaker, or by moving the flaps and flap lever to the retracted position.

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10.5. Leading Edges Slats. 10.5.1. General. Primary Control. The leading edge slats are controlled by inputs from the flap lever. The flap lever operates a cable system to the slat aft quadrant which is connected to the inboard and outboard slat power drive units (PDU). A hydraulic motor powers each power drive unit (PDU) gearbox, which operates drive shafts, gearboxes and rotary actuators connected to the slats. There is a separate PDU and drive system for the inboard and outboard slats. Krueger seal flaps, operated by the inboard slat drive system, are fully extended when the inboard slats are in takeoff (sealed) position. Slat Alternate Control. The PDU gearbox can also be operated by an electric motor for alternate system operation. The electric motor is controlled by the flap slat electronic unit (FSEU). CAUTION : BEFORE OPERATING FLAPS OR SLATS, ENSURE THAT ENGINE STRUT ACCESS DOORS. INBOARD FAN COWLING, AND THRUST REVERSER COWLING ARE NOT IN THE PATH OF SLATS, TO PREVENT DAMAGE.

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L.E. SLATS - GENERAL EFFECTIVITY ALL

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10.5.2. INB / OUTB Slats. Inboard Slat Drive and Tracks. Each inboard slat is extended and retracted by two control rods each connected to a rotary actuator arm and a fitting on the aft side of the slat. The main support for the slat is by an “A” frame at the inboard end, a support arm at the center and a main track, with emergency down stop, at the outboard end, all mounted to the wing front spar. The slat is held in position on the main support components and the control rods by three auxiliary track arms. The slat is attached to each auxiliary track arm at two places to prevent slat rotation and provide slat angle positioning by the profiled auxiliary track. The auxiliary tracks position the inboard slats at 12° when extended to the intermediate or takeoff position and at 30.3° when fully extended. Outboard Slat Drive and Tracks. Each outboard slat is extended and retracted by two control rods each connected to a rotary actuator arm and a fitting on the aft side of the slat. Two main tracks, with emergency down stops, support the slat. These tracks extend into recesses in the wing fuel tanks when the slats are retracted. The slat is held in position on the main tracks and control rods by two auxiliary track arms. The slat is attached to the auxiliary track arms at two places to prevent slat rotation and, provide slat angle positioning by the profiled auxiliary track. The auxiliary tracks position the outboard slats at 26° when extended to the intermediate (takeoff) position and at 35° when fully extended.

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OUTBOARD SLAT DRIVE & TRACKS EFFECTIVITY ALL

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10.5.3. Slat Position. Sensor. A proximity sensor is installed on wing leading edge structure near the inboard and outboard auxiliary tracks-on each slat. The sensor is actuated by either a retract target, riveted on the auxiliary track arm, or a roller bolt assembly at the aft end of the auxiliary track arm that holds the track arm on the track. On each slat one sensor is installed farther aft on wing structure than the sensor near the other auxiliary track with its retract target installed the same distance aft on the auxiliary track arm. Operation. When the slat is fully retracted the retract target on the auxiliary track arm is near to the proximity sensor. When the slat is in the intermediate (takeoff) position the auxiliary track roller ball assembly target is near on the inboard auxiliary track on the outboard slats and the outboard auxiliary track on the inboard slats. When the slat is fully extended, the near sensor/targets are reversed, with the roller ball assembly target near on the outboard auxiliary track for the outboard slats and the inboard auxiliary track for the inboard slats. Maintenance Practices. A target out of adjustment or a faulted sensor would be detected by the PSEU as a slat asymmetry condition. The PSEU built-in test (BITE) would identify the faulted component by a sensor/ target code and sensor or target fault light.

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10.5.4. LE Slat Alternate. The slat inboard and outboard power drive unit (PDU) alternate drive electric motors are controlled by the flap slat electronic unit (PSEU). The FSEU receives command inputs from the alternate flaps position selector switch and slat drive position from rotary variable differential transformers (RVDT) on each PDU. Using these inputs, section 3 of the FSEU controls relays to provide power to the reversible electric motors on both PDUs. The slat alternate arm switch directly controls the bypass valves on the PDUs during alternate drive operation to remove the hydraulic motor lock on the PDU gearbox. The arm switch also controls power to the FSEU for relay control. The FSEU compares the command and slat drive position inputs separately in the inboard and outboard slat systems for individual, closed-loop motor control. Electric motor turn off does not occur simultaneously on the inboard and outboard slat systems due to different degrees of extension at the intermediate (takeoff) and fully extended positions.

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LE SLAT ALTERNATE DRIVE CONTROL EFFECTIVITY ALL

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10.5.5. Slat Asymmetry and Failure Protection Shutdown. Asymmetry Protection Shutdown : - An asymmetry condition detected by the proximity switch electronic unit (PSEU), in the inboard or outboard slat systems, is signaled to the flap slat electronic unit (FSEU) for protective shutdown. The asymmetry signal is latched in the FSEU and annunciated until the signal is removed by the PSEU. If the flap lever is moved with a latched asymmetry, the FSEU energizes the inboard or outboard slat fail protection and asymmetry relay to bypass the hydraulic motor. The shutdown and asymmetry annunciation are latched and will not reset if the PSEU removes the asymmetry signal. Cycling the alternate slat arm switch on and off will reset the asymmetry shutdown latch, de-energize the fail relay, and energize the bypass valve normal relay to restore hydraulic motor operation and clear the asymmetry annunciation. Failure Protection Shutdown : - The FSEU compares the flap lever and power drive unit (PDU) transmitters (RVDTs) to detect uncommanded movement of the inboard or outboard slat systems. If the slats are moving away from the flap lever position, the FSEU operates the fail relay to bypass the hydraulic motor. The failure shutdown latch can be reset by cycling the alternate slat arm switch to de-energize the fail relay, energize the normal relay and restore hydraulic motor operation. The latch also resets when the flap lever is up and the slats are retracted.

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10.5.6. Slat Disagree. The proximity switch electronic unit (PSEU) compares slat position signals from proximity sensors at each slat with flap lever or alternate flap position selector switch position inputs from the flap slat electronic unit (FSEU) to detect a slat disagreement. Operation. When the flap lever is moved during primary drive slat system operation or the alternate flap position selector switch is moved during alternate operation, the PSEU sends a disagree signal to the FSEU. When the proximity sensors show agreement with slat command the PSEU removes the disagree signal. A PSEU disagree input is inhibited in the FSEU when any of the following conditions exist. - Either inboard or outboard slat PDU is moving toward the flap lever or alternate flap position selector switch position (command) in primary or alternate operation. - The flap drive is moving toward the flap lever position in primary operation. - The slats are commanded up in primary or alternate operation and the flaps are not retracted. - A slat asymmetry condition exists. If the PSEU disagree signal is present after the inhibits are removed, the FSEU causes display of the LEADING EDGE light and LE SLAT DISAGREE caution message after ten seconds.

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11. WARNINGS. 11.1. Stall Warning System 11.1.1. General. The stall warning system has two digital stall warning computers (SWC) whose function is to calculate when the airplane is nearing a stall condition and provide a warning through operation of the stick shakers and stick nudger. Another function of the SWC’s is to input to the windshear detection and guidance systems for visual and aural warning annunciation and flight instrument display.

11.1.2. System Description. Each SWC operates a separate stick shaker. The output of both SWC’s is required for stick nudger operation. The SWC’s input to the ground proximity warning computer (GPWC) and the electronic flight instrument system (EFIS) for the windshear detection and guidance system and to the EICAS computers for fault annunciation. Inputs to the SWC’s are flap position (flap/stab position module - FSPM), slat movement (flap/slat electronic unit - FSEU), slat position (proximity switch electronic unit - PSEU), speedbrakes down or not down (spoiler control module - SCM) body pitch angle and rate (inertial reference system -IRS), dual power sources (PSM), mach, true airspeed, computed airspeed, indicated angle of attack (air data computers) and air ground sensing (air/ground relays controlled by the PSEU). Each SWC has a test switch and a BITE display.

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11.1.3. Components Description. Stick Shakers. Captain’s and first officer’s stick shakers are located on the elevator torque tube sections under the flight compartment floor. Shakers are accessible from the access door forward of the nose wheel well. Stick Nudger. A single stick nudger actuator, located on the elevator feel and centering unit provides a column forward force by repositioning the feel and centering unit input levers. Stall Warning Computers The left and right stall warning computers are located in the warning electronics unit (P51 panel) accessible from the main equipment area.

B767/27/101 Flight controls

Interfacing System Components Center and right flap/stabilizer position modules (P50). Proximity switch electronic unit (E1-2). Flap/slat electronic unit (E2-4). Air data computers (E1-3). Inertial reference system (E1-6) Spoiler control modules 2L and 1R (E1-1), (E2-1). Air ground relays (controlled by the PSEU) (P36/P37 panels).

Test Panel Left and right stall warning test switches are located on the P61 side panel. Each switch is a momentary toggle switch for testing a stall warning system.

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11.2. Takeoff Configuration Warning. 11.2.1. General. The system warns pilots of improper airplane configuration for takeoff. Input signals, are processed by the takeoff configuration warning module. Airplane in improper T/O configuration is indicated by : - Two-tone siren aural from aural warning speakers, - Master warning lights on P7 glareshield, - CONFIG warning light on P1-3 panel, - Warning message(s) displayed on EICAS display unit.

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11.2.2. Components Description. The takeoff configuration warning system monitors the configuration of certain critical systems and advises the pilot of improper airplane configuration prior to takeoff. Control. Improper airplane configuration detected when either throttle is advanced (engine speed card) and is announced by : - Illumination of the red, CONFIG light on the P1-3 panel, - Illumination of the red master warning lights on the P7 glareshield panel, - Actuation of the aural warning sirens, - Annunciation of the appropriate fault message on the EICAS display. Input. Flap/slat disagreement from the flap slat electronic unit (FSEU). LE slat position from the proximity switch electronic unit (PSEU). TE flap position & stabilizer position from the center flap/stab position module (FSPM) Speedbrake position from the S493 speed brake switch. Parking brake position from the S459 parking brake switch. Air/gnd mode from the PSEU. Self Test Check the T/O warning system by advancing either throttle (ENGINE SPEED card from EICAS thrust discrete) or by actuating T/O configuration test switch on P61 side panel with one of the monitored systems not in T/O configuration.

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11.3. Landing Configuration Warning. 11.3.1. General. It warns the pilots when the airplane is in improper configuration for landing. Input. Input signals, supplied from airplane sensors, avionics systems and pilots, are processed by the landing configuration warning module. The landing configuration warnings are generated as a result of the following conditions : - Radio altitude of less than 800 feet, either throttle at idle and the landing gear not down and locked, - Flaps in the landing range and the landing gear not down and locked, - Radio altitude of less than 800 feet and spoilers deployed. Annunciations. Airplane in improper landing configuration is indicated by : - Two-tone siren aural from aural warning speakers, - Master warning lights on P7 glareshield, - CONFIG warning light on P1-3 panel, - Warning message GEAR NOT DOWN displayed on EICAS upper display unit, - SPEED BRAKE caution light on P1-3 panel.

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11.3.2. Components Description. The purpose of the landing configuration warning system is to warn the flight crew of a gear not down and locked condition when the aircraft is configured for landing. Configuration. Improper configuration of the airplane for landing is indicated by : - Level A warning and GEAR NOT DOWN message. Improper use of speedbrake below 800 feet radio altitude is indicated by : - Level B caution and SPEED BRAKE EXT message. Conditions Monitored. Landing gear position from the PSEU. Slat position from the PSEU. Flap position from the center flap/stab position module. Altitude from the radio altimeter in the E5 rack. Thrust lever position from the autothrottle switch pack Speedbrake position from a spoiler control module (speedbrake LVDT). Signal Processing. The input signals are processed by the landing configuration warning module located in the warning electronic unit. The module may be tested by use of the configuration test switch in the landing position.

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11.4. Speedbrake Warning. The purpose is to provide an aural and visual caution if the speedbrakes are deployed when the airplane is between 800 feet and 15 feet altitude, or if the speedbrakes are deployed when the aircraft is above 15 feet and the trailing edge flaps are in the landing range. Caution Conditions. Whenever the speedbrake handle is advanced beyond the ARM position and the airplane is more than 15 feet above the ground, caution annunciations are initiated by the landing configuration warning module. The annunciations occur under the following conditions : - Condition 1 : Speedbrake handle beyond ARM position, radio altitude greater than 15 feet and flaps in the landing position (25 or 30 units). In this case, the speedbrake drag would tend to cancel much of the extra lift created by the flaps being in the landing range. - Condition 2 : Speedbrake handle beyond ARM position and the radio altitude greater than 15 feet, but less than 800 feet.

Test Results. During a speed brake test the microprocessor in the landing configuration warning module performs an internal test to check itself and the left radio altimeter’s input data bus for validity. If the test results are correct, gates 3 and 4 are enabled. This causes the SPEED BRAKE EXT EICAS message to be displayed and the SPEED BRAKE caution light to illuminate. The master CAUTION lights and level B caution aural are inhibited by the EICAS during a ground test. However, these two annunciations would occur during an in light test. NOTE : Since other landing warning configuration module functions are also checked during this test, the following warning annunciations will also be generated, as previously described : - GEAR-NOT-DOWN EICAS message (red), - CONFIG light (red) illuminated on P1-3, - Master WARNING lights (red) illuminated on P7, - Siren sounds.

Test Initiation. A speedbrake test function can be initiated in the air or on the ground by advancing the speedbrake handle past the ARM position and then actuating the LDG CONFIG test switch on the P61 test panel. The test annunciations remain as long as these test conditions remain.

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