CF6-80C2 ATA 70-80 Power Plant LLTT.pdf

Training Manual BOEING 767--300 POWER PLANT (GE / CF6 -- 80C2) ATA 71 -- 80

For training purpose and internal use only. Copyright by Lufthansa LAN Technical Training S.A. All rights reserved. No parts of this training manual may be sold or reproduced in any form without permission of:

Lufthansa LAN Technical Training S.A. Clasificador 74 Av. Américo Vespucio 901, Renca Santiago -- Chile Tel. +56 (0)2 601 99 11 Fax +56 (0)2 601 99 24 www.lltt.cl

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

ATA -- 71 POWER PLANT TABLE OF CONTENT General Data Cowlings Inlet Cowl Fan Cowl Fan Cowl Latch Ajustment Fan Cowl Chine Thrust Reverser Thrust Reverser Latch Thrust Reverser Latch Ring Thrust Reverser Opening Actuator Thrust Reverser Deflection Limiter Core Cowl Core Cowl Latch Adjustment Turbine Exhaust and Plug Engine Mounts Fwd Engine Mount Aft Engine Mount Engine Vent and Drains Engine Hazard Areas Engine Entry Corridor Engine Noise Hazard Areas

002 004 006 010 014 016 018 020 022 024 032 034 038 040 042 043 045 046 048 052 054

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ENGINE GENERAL

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POWER PLANT GENERAL The General Electric CF6--80C2F is a high bypass ratio, axial flow, dual--rotor turbofan engine. The two strut--mounted engines supply airplane thrust, and power the electrical, pneumatic and hydraulic Systems. The engine data and engine assembly identification plates are attached to the left tan case. Engine specifications are listed on the next graphic.

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ENGINE GENERAL

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ENGINE GENERAL

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Engine Data Page: 5

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ENGINE GENERAL

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ENGINE COWLING Purpose The cowling is an aerodynamically smooth protective cover surrounding the engine, engine--mounted components, and accessories. The cowling directs airflow around and through the engine. Description The cowling for each engine includes the inlet cowl, fan cowl panels, thrust reverser halves and core cowl panels. There are access doors and openings on the cowling for maintenance, servicing and pressure relief. An exhaust sleeve and exhaust plug direct the hot turbine exhaust gases exiting the low--pressure turbine. Hinges hold the fan cowl panels, thrust reversers and core cowls to the strut. The inlet cowl, exhaust sleeve and exhaust plug are bolted directly to the engine. An aerodynamic chine is mounted on the inboard fan cowl panel.

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Cowl Opening Secuence Open the fan cowl panels first, then the thrust reverser, then the core cowl , panels. Close the core cowl panels first, then the thrust reverser, then the fan cowl panels.

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ENGINE GENERAL

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Engine Cowling Page: 7

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INLET COWL Purpose The inlet cowl directs air into the fan. It is mounted on the engine fan case forward flange. Description The inlet cowl has an inner barrel, an outer barrel, an inlet lip, and forward and aft bulkheads. It is an aluminum structure with Kevlar--graphite external panels. Honeycomb acoustic panels line the inner surface of the inlet cowl to reduce air noise. Thermal bleed air prevents ice from forming on the inlet cowl leading edge. An anti--ice air exhaust port is located on the aft. bottom of the cowl. There are provisions for a service interphone jack on the lower left side. (Not operational on the Boeing 767). There are four hoist points on the outer barrel for attaching a sling to remove and install the cowl.

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ENGINE GENERAL

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ENGINE GENERAL

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Inlet Cowl Page: 9

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

INLET COWL REMOVAL AND INSTALLATION General Remove the fan cowl panels before removing the inlet cowl. The inlet cowl weighs about 527 pounds (239 Kg). DURING INLET COWL REMOVAL / INSTALLATION, DO NOT LEAVE TOOLS OR OTHER OBJECTS IN AIR INLET. FOREIGN OBJECTS CAN CAUSE SEVERE DAMAGE TO ENGINE WHEN INGESTED The TAl duct must be disconnected. CAUTION:

ADJUST SLING TO TAKE ONLY THE WEIGHT OF THE INLET COWL. ADDITIONAL WEIGHT CAN DAMAGE COWL AND SLING. A crane and sling assembly is used to remove the inlet cowl. After the mount bolts are removed, pull the cowl forward to clear the index pins. CAUTION:

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Installation Make sure that the index pins are installed on the inlet cowl. Align the cowl with the index pin receptacles on the engine flange. Install the mount bolts. Connect the thermal anti--ice duct, and install the fan cowl panels.

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For Training Purposes Only

ENGINE GENERAL

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Inlet Cowl Removal and Installation Page: 11

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

FAN COWL PANELS General The fan cowl panels are hinged to the strut and align with the inlet cowl and thrust reverser. Panels are latched together at the bottom centerline with three flush--mounted tension latches. The fan cowl panels open for access to components on the engine fan case. Each fan cowl overlaps the corresponding thrust reverser half. The right fan cowl panel has an access door to service the engine oil tank without opening the fan cowl. This panel is also a pressure relief panel. There are two hold--open rods on each fan cowl panel. The hold--open rods engage brackets on the tan case and extend to hold the tan cowl open in either of two positions. The tree ends of the rods are stowed in receivers on the cowl. Opening Fan Cowl Panels Engage the forward hold--open rod first, then engage the aft hold--open rod.

Closing Fan Cowl Panels Close the corresponding thrust reverser half before closing the fan cowl panel. Disengage the aft hold open rod first, then disengage the forward hold open rod. Retract the sleeve on the hold open rod and disengage the rod from the engine mounted receiver. Release the secondary lock and slide the outer collar to unlock the hold open rod. The UNLOCKED indication is then visible. Repeat the unlock procedure for the inner collar. Retract the hold--open rod and engage it into the fan cowl panel receiver. DO NOT ALLOW FAN COWL PANEL TO SLAM CLOSED. DAMAGE TO FAN COWL PANEL AND / OR ENGINE COMPONENTS MAY RESULT. Push the fan cowl panels together and engage the latches. CAUTION:

ADEQUATE SUPPORT OF FAN COWL PANEL MUST BE MAINTAINED WHILE ENGAGING HOLD OPEN RODS TO PREVENT INJURY TO PERSONNEL AND / OR ENGINE COMPONENTS. Retract the sleeve at the receiver end of the hold open rod to remove the rod from the receiver. Fully extend and lock the outer rod segment. Push in on the secondary lock and pull back the inner collar to unlock the inner segment. Fully extend and lock the inner segment. Check that the red UNLOCKED bands at the collars are not visible WARNING:

ENSURE THAT HOLD OPEN ROD IS FULLY EXTENDED AND LOCKED TO PREVENT ACCIDENTAL CLOSING OF COWL PANEL. PERSONNEL STRUCK BY FALLING COWL PANEL COULD BE SERIOUSLY INJURED. ROD IS NOT LOCKED IF RED BAND WITH THE WORD ’UNLOCKED” IS VISIBLE. IF RED BAND IS VISIBLE, ROD WILL RETRACT UNDER LOAD. Hold the sleeve in, engage hold open rod into the engine--mounted receiver and release the sleeve.

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WARNING:

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Fan Cowl Panels Page: 13

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

FAN COWL REMOVAL AND INSTALLATION Removal Open the fan cowl panel to be removed When removing the ball lock pins, check that the cowl panel hinge fittings rest on the roll pins. ADEQUATELY SUPPORT FAN COWL PANEL DURING HANDLING. FAN COWL PANELS WEIGH ABOUT 110 POUNDS EACH. Manually support the fan cowl panel and disengage the hold open rods. Use the three lift sling attach points to lift the fan cowl outward from the roll pins. WARNING:

CAUTION:

RAISING OR LOWERING FAN COWL PANEL AFTER REMOVAL OF HINGE BALL LOCK PINS MAY DAMAGE UPPER COWL SEAL. CAREFULLY LIFT PANEL OUTWARD FROM STRUT HINGE FITTING TO AVOID DAMAGE TO SEAL.

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Installation Position the fan cowl panel hinge fittings on the roll pins at each hinge location. Rotate the panel 55 ° open to align the hinge fitting holes. Install the ball lock pins and cotter pins. Adjust the fan cowl latches.

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ENGINE GENERAL

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Fan Cowl Panel Removal and Installation Page: 15

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

FAN COWL PANEL LATCH ADJUSTMENT Adjustment -- Latches and Shims Adjusting the fan cowl panel latches is necessary for panel security and aerodynamic smoothness. Adjust the latches whenever either fan cowl panel or thrust reverser half is replaced. DO NOT USE OVER 100 POUNDS FORCE TO PUSH LATCH HANDLE CLOSED. EXCESSIVE FORCE CAN DAMAGE LATCH. Close the fan cowl panels, using hand--pressure, and close the latches. An adjustment is required if the cap between left and right fan cowl panels is not between .06 and .18 inches. The adjustment is made with shims. CAUTION:

Test -- Force Required to Close Latches CAUTION:

DO NOT USE OVER 100 POUNDS FORCE TO PUSH LATCH HANDLE CLOSED. EXCESSIVE FORCE CAN DAMAGE LATCH.

DO NOT ROTATE KEEPER EYE BOLT TO ADJUST LATCH TENSION. DAMAGE TO KEEPER MAY RESULT. If the force required to close the latch is not between 50 and 100 pounds, open the latch handle to release tension on the keeper. Insert a hex wrench into the adjustment star within the keeper mounting and rotate the adjustment star with the hex wrench. The latch keeper mounting shows the direction to rotate the adjustment star to increase the load. Properly adjusted latches close with a loud pop. Close the fan cowl latches and check that all the latch handles are even with the fan cowl panel.

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CAUTION:

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ENGINE GENERAL

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Fan Cowl Panel Latch Adjustment Page: 17

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FAN COWL CHINE The fan cowl chine improves airplane aerodynamic characteristics at low air speeds. The chine is installed at 45 ° from the fan cowl panel top centerline on the inboard fan cowl panels. A fiberglass insulator is mounted between chine and fan cowl panel.

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ENGINE GENERAL

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ENGINE GENERAL

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Fan Cowl Chine Page: 19

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

THRUST REVERSER General When the thrust reverser is stowed, it acts as a cowl for efficient thrust. When the thrust reverser is deployed, fan exhaust air is deflected forward to slow down the airplane. The thrust reverser halves are attached to the strut and align with the fan cowl and core cowl. Opening the thrust reverser permits access to components on the high pressure compressor case and accessory gearbox. Each thrust reverser half overlaps the corresponding core cowl panel. They are mounted to the lower part of the strut with three hinges. The thrust reverser halves are latched closed with tension latches and the thrust reverser latch ring assembly. The thrust reverser latch ring assembly has upper and lower latches, upper and lower latch handles and upper latch cable. Major components for the thrust reverser system are mounted to the reverser torque box and fixed structure.

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Operation The inner and outer duct walls make a flow path for fan air exhaust. Translating cowls, drag links and blocker doors direct fan exhaust through the deflectors when the thrust reverser is deployed. Pneumatically powered center drive units and ball screw actuators move the translating cowls. The deflectors are covered by the translating cowl, when stowed. The translating cowl is lined with acoustical material to reduce noise. The deflectors are also called cascade segments or cascade vane segments.

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ENGINE GENERAL

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Thrust Reverser Page: 21

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

THRUST REVERSER TENSION LATCHES General The thrust reverser halves are latched together by three tension latches along the bottom split--line. The latches are mounted within the area covered by the access and blow out doors on the bottom of the thrust reverser. The forward blow out door must be opened first and closed last. Latch hooks are on the left half and fit over latch pins on the right half. Latch tension is adjustable.

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Adjustment The fan cowl panels must be open. The access and blow out doors must be open. Unlatch all three tension latches in order, starting with the aft latch, working forward. Check the tension latches for damage. The tension latch handle closing force is measured with a spring scale. Adjust tension latches from forward to rear. Adjust the closing force by loosening the latch bolt nut and rotating an octagonal offset bushing.

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ENGINE GENERAL

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Thrust Reverser Tension Latches Page: 23

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

THRUST REVERSER LATCH RING ASSEMBLY General The thrust reverser latch ring assembly secures the outer leading edge of the thrust reverser halves to the aft flange of the fan stator case. It transmits reverser loads into the engine fan frame instead of the strut hinges. This assembly is mounted around the leading edge of each thrust reverser half. Access is through the fan cowl panel. The upper latch of the mounting ring is a hook that engages a U bolt on top of the stator case. The U bolt is adjustable to control upper latching force. The bottom latch is a barrel nut that fits into a claw type clevis bracket at the bottom of the fan case. The barrel nut is adjustable to control lower latching force. Upper and lower latch handles open and close upper and lower latches. The upper latch cable is adjustable. The thrust latch ring assembly is removed by removing the attachment bolts (not shown).

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Operations To open the thrust reverser latch ring assembly, pull the lower latch handle outward until the latch pin bottoms in the slot. Rotate the upper latch handle outward disengaging the latch pin from the slot. The upper latch is now disengaged from the U bolt. Rotate the lower latch handle outward, disengaging the barrel nut from the clevis bracket. To close the thrust reverser ring latch assembly, engage the barrel nut with the clevis and rotate the lower latch handle inward rotate the upper latch handle inward engaging the latch pin in the slot. The upper latch engages the U bolt.

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ENGINE GENERAL

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Thrust Reverser Latch Ring Assy Page: 25

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THRUST REVERSER OPENING ACTUATOR The thrust reverser opening actuator permits each thrust reverser half to be opened with a portable hydraulic pump. Each thrust reverser opening actuator is mounted to a bracket on each side of the airplane strut. The thrust reverser opening relief valve is mounted to the multiple connector. A flexible hose is connected from the strut T-- fitting to the thrust reverser opening actuator inlet fitting. The inlet fitting has a restrictor to limit the rate of closure. If a hydraulic line ruptures, or if the thrust reverser half is closing too fast, the restrictor ensures that the thrust reverser half takes at least 15 seconds to close. A 25 micron filter at the input fitting protects the restrictor and actuator assembly from fluid contamination. The thrust reverser opening relief valve relieves high system pressure and is set to open at 4350 -- 4500 psig.

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ENGINE GENERAL

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ENGINE GENERAL

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

THRUST REVERSER OPENING / CLOSING Operation Thrust Reverser Opening WARNING:

USE THE THRUST REVERSER HYDRAULIC POWER OPENING SYSTEM ONLY FOR OPENING AND CLOSING THE REVERSER HALVES.THE SYSTEM SHOULD NEVER BE USED AS A HOLD OPEN DEVICE. ALWAYS SECURE EACH OPENED REVERSER HALF WITH A HOLD OPEN ROD, TO PREVENT SERIOUS INJURY DUE TO ACCIDENTAL OR INADVERTENT CLOSURE. KEEP ALL PERSONNEL CLEAR OF AREAS UNDER AND BETWEEN REVERSER HALVES DURING OPENING AND CLOSING CYCLES.

CAUTION:

BE SURE LEADING EDGE SLATS ARE RETRACTED AND LOCKED BEFORE OPENING THRUST REVERSER. FAILURE TO DO SO MAY RESULT IN DAMAGE TO THRUST REVERSER, LEADING EDGE SLATS AND / OR WING.

CAUTION:

DO NOT OPEN THRUST REVERSER BEYOND THE 20 DEGREES POSITION WITH THE THRUST REVERSER TRANSLATING COWLS EXTENDED. DAMAGE TO TRANSLATING COWLS OR STRUT MAY RESULT.

INSTALL HOLD OPEN ROD BALL--LOCK PIN WITH PLUNGER BUTTON UP. Close the pump valve and operate the pump. The fluid is pumped into the thrust reverser opening actuator. Open the reverser half far enough to connect the hold--open rod to the fan case support.For the 20 degree position, connect the hold open rod without extending the rod. For the 45 degree position, remove the ball lock pin and extend the rod to its full length, then install the ball lock pin. Install the ball lock pin with plunger button up. Release the hydraulic pressure by slowly opening the hydraulic pump valve. Disconnect the pump hose and install the dust cover. CAUTION:

ENSURE THAT LATCH RING UPPER LATCH HANDLE IS FULLY OVER--CENTER BEFORE OPENING REVERSER. FAILURE TO TOTALLY DISENGAGE UPPER LATCH HOOK FROM U BOLT COULD RESULT IN DAMAGE TO EQUIPMENT. The fan cowl panel must be opened and secured before the corresponding thrust reverser half is opened. Open the blowout and access doors. Release the thrust reverser lower tension latches. Release the thrust ring latch assembly by rotating the upper and lower latch handles. Attach a hose from hydraulic hand pump to the quick--disconnect hydraulic connector. The connectors are located on the aft fan case (5:00 for the right thrust reverser half and 7:00 for the left thrust reverser half).

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CAUTION:

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ENGINE GENERAL

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Thrust Reverser Opening / Closing Page: 29

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

Thrust Reverser Closing The core cowl panel must be latched closed before the corresponding thrust reverser half is closed. CAUTION:

ENSURE CORE COWL PANEL IS FULLY CLOSED WHEN CLOSING THRUST REVERSER HALF OR DAMAGE TO CORE COWL MAY OCCUR.

WARNING:

DO NOT STAND BETWEEN ENGINE AND THRUST REVERSER WHEN CLOSING THRUST REVERSER. INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT COULD OCCUR.

CAUTION:

OBSERVE THAT THE VEE--FLANGE GUIDES INTO ENGINE VEE GROOVE AND THAT FULL ENGAGEMENT IS TAKING PLACE WHEN CLOSING THRUST REVERSER. DAMAGE TO THRUST REVERSER MAY RESULT FROM MISALIGNMENT.

ENSURE LATCH RING UPPER LATCH HANDLE IS IN THE FULLY OPEN POSITION BEFORE CLOSING THRUST REVERSER. DAMAGE TO UPPER LATCH U--BOLT SPRING RETAINER MAY RESULT. Remove the dust cover from the hydraulic connector and connect the hydraulic pump hose. Close the pump valve and operate the pump until the reverser weight is removed from the hold--open rod. Stow the hold--open rod. Slowly open the pump valve to close the reverser. CAUTION:

NOTE:

WITH THE PUMP VALVE OPEN, THE REVERSER SMOOTHLY CLOSES IN APPROXIMATELY 15 SECONDS.

WHEN SECURING THRUST REVERSER, VERIFY THE UPPER LATCH HOOK HAS ENGAGED THE U--BOLT Secure the thrust ring latch assembly and the three tension latches. Disconnect the pump hose. Close access door, then close the blowout door.

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CAUTION:

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ENGINE GENERAL

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Thrust Reverser Opening / Closing Page: 31

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

THRUST REVERSER REMOVAL AND INSTALLATION Removal To remove the thrust reverser, the fan cowl panel and skirt fairing must first be removed, applicable circuit breakers opened, leading edge flaps retracted and the thrust reverser deactivated. Connect the thrust reverser sling to the thrust reverser at the four attach points. Support the thrust reverser with a lifting device. Remove the three hinge bolts. Disconnect the pneumatic supply line, sense line, and electrical connector. WARNING:

ENSURE THRUST REVERSER IS SUPPORTED SECURELY BY THE SLING HOIST AND HOLD--OPEN RODS. THRUST REVERSER COULD CLOSE SUDDENLY CAUSING SEVERE INJURY TO PERSONNEL, AND / OR DAMAGE TO EQUIPMENT.

CAUTION:

THE THRUST REVERSER HALVES CAN NOT BE LIFTED OR MOVED UNLESS ALL 16 CASCADE VANE SEGMENTS ARE INSTALLED. DAMAGE TO THRUST REVERSER STRUCTURE MAY RESULT. CONTROL SWING OF THRUST REVERSER WITH TAG LINES TO PREVENT THRUST REVERSER SWINGING INTO ENGINE OR EQUIPMENT.

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Installation Align the thrust reverser with the three hinge fittings and install the hinge bolts. Check that the clearance is correct between the thrust reverser fitting and strut fitting. Lower the thrust reverser and remove the lifting device and sling. When raising the thrust reverser for installation, it is suspended at an angle of 45 degrees.

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ENGINE GENERAL

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Thrust Reverser Removal and Installation Page: 33

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

THRUST REVERSER DEFLECTION LIMITER ADJUSTMENT General The deflection limiter is a pad that makes sure compression of the fire and drain seals is correct. The pads also control the clearance between the two thrust reverser halves and between the thrust reverser halves and the engine strut. There are three deflection limiters on the left and right thrust reverser upper bifurcations. There are two on the right lower bifurcation The deflection limiter must be adjusted after a thrust reverser is removed and replaced. Adjustment Procedure

CAUTION:

ENSURE ACCESS PANEL DOOR IS CLOSED AND LATCHED BEFORE CLOSING BLOWOUT DOOR. WITH DOORS CLOSED, MAKE SURE DOOR RETENTION PINS ARE ENGAGED. PRELOAD MUST NOT EXIST ON BLOWOUT DOOR LATCHES WITH DOOR CLOSED. BLOWOUT DOOR RETENTION CABLES MUST BE PROPERLY STOWED TO AVOID PRELOAD OR INTERFERENCE.

FAILURE TO DEACTIVATE THRUST REVERSER HALVES FOR GROUND MAINTENANCE COULD RESULT IN INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT. Deactivate both thrust reverser halves before working on the engine. This procedure uses petroleum jelly as a parting agent on the three upper bifurcation deflection limiter wear pads on each side of the strut, and on the two lower pads on the left reverser half. Modeling clay and petroleum jelly or transfer dye is used to measure the contact between the strut and thrust reverser. Apply the petroleum jelly or dye to the strut along the fire seal contact area. Apply clay to the upper bifurcation deflection limiters on each reverser half and the two lower bifurcation deflection limiters on the right thrust reverser half. Close, latch, and then open the reversers. The resulting depression on the clay, and the transfer of the dye or jelly on the fire seals, tells if the deflection limiters are adjusted properly. Add or remove shims if adjustment is necessary. Also check tension latch closing force and access / blowout door overlap at this time, and adjust as necessary.

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WARNING:

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ENGINE GENERAL

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Thrust Reverser Deflection Limiter Adjustment Page: 35

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

CORE COWL PANELS General The left and right core panels cover the turbine case section of the engine. They open to allow access to the combustion and turbine cases of the engine. The core cowl panels are attached to the strut with hinges, align with the inner barrel of the thrust reverser on the forward edge, and rest against the engine exhaust sleeve on the aft edge. The panels are latched together with three flush--mounted tension latches at the bottom. A hinged pressure relief door with a latch is installed on the right core cowl panel. Two lanyards restrain the door when it is open. Fire shields are located inside the panels. A hold--open rod on each cowl is extended and connected to a bracket on the engine to hold the cowl 50° open. When the rod is not in use, the free end is stowed in a receiver on the cowl. Opening Core Cowl Panels Open the fan cowl panels and thrust reverser halves before opening the core cowl panels. BE SURE FAN COWL PANELS ARE OPENED AS REQUIRED BY 71--11--06 BEFORE OPENING THRUST REVERSER. FAILURE TO FOLLOW 71--11--06 COULD RESULT IN INJURY TO PERSONNEL AND / OR DAMAGE TO FAN COWL PANELS, CORE COWL PANELS, AND THRUST REVERSER. Release the core cowl latches and disengage the hold--open rods from the receivers. Fully extend the rod to the locked position. The red UNLOCKED indicator band must not be visible. WARNING:

Closing Core Cowl Panels ADEQUATE SUPPORT OF CORE COWL PANEL MUST BE MAINTAINED WHILE HOLD--OPEN RODS ARE BEING DISENGAGED TO PREVENT INJURY TO PERSONNEL AND / OR ENGINE COMPONENTS. Retract the sleeve at the receiver end of the hold--open rod to disengage the rod. To unlock the hold--open rod from its extended position, rotate and slide the collar in the direction indicated and push the secondary lock. The hold open rod is now retracted allowing the collar to move to its original position. The UNLOCKED indication is visible. Connect the hold--open rod to the receiver on the cowl to stow it. WARNING:

DO NOT ALLOW CORE COWL PANELS TO SLAM CLOSED. DAMAGE TO PANEL AND / OR ENGINE COMPONENTS MAY RESULT. Stow the hold--open rod and lower the core cowl panel. CAUTION:

ENSURE THAT HOLD--OPEN ROD IS FULLY EXTENDED AND LOCKED TO PREVENT ACCIDENTAL CLOSING OF COWL PANEL. PERSONNEL STRUCK BY FALLING COWL PANEL COULD BE SERIOUSLY INJURED. ROD IS NOT LOCKED IF RED BAND WITH THE WORD UNLOCKED IS VISIBLE. IF RED BAND IS VISIBLE, ROD WILL RETRACT UNDER LOAD. Hold the sleeve retracted to engage the hold--open rod to the engine--mounted bracket.

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CAUTION:

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ENGINE GENERAL

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Core Cowl Panels Page: 37

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

CORE COWL PANEL REMOVAL AND INSTALLATION Removal Open the core cowl panel to be removed. ADEQUATELY SUPPORT CORE COWL PANEL DURING HANDLING. RIGHT CORE COWL PANEL WEIGHS ABOUT 90 POUNDS. LEFT CORE COWL PANEL WEIGHS ABOUT 65 POUNDS. Manually support the core cowl panel using the sling attach points. Stow the hold--open rod. Remove the ball lock pin from each hinge fitting and lift the panel off the hinge fittings. WARNING:

For Training Purposes Only

Installation Position the core cowl panel on the strut and align with the hinge fitting holes. Install ball lock pins and cotter pins at each hinge location. Close the core cowl panel. Adjust the latches if necessary.

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Figure 18 SCL

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Core Cowl Panel Removal and Installation Page: 39

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

CORE COWL PANEL LATCH ADJUSTMENT Adjustment -- Latches and Shims The core cowl panel latches are adjusted for panel security and aerodynamic smoothness. Adjust the latch whenever either thrust reverser half or core cowl panel is replaced. FAILURE TO PROPERLY ADJUST LATCHES AND SHIMS MAY ALLOW LATCHES TO DISENGAGE IN FLIGHT RESULTING IN LOSS OF COWL. With the core cowl panels open, see that the keeper eye bolts do not rotate, and that the retention pins are not sheared off. If the keeper eye bolt rotates, replace the broken or damaged keepers and / or latches immediately. With the core cowl panels closed and latched, measure the gap between the core cowl panels at each latch. Adjust the gap if it is greater than .220 inch, using shims and a bearing pad. CAUTION:

Test -- Force Required to Close Latches DO NOT USE OVER 100 POUNDS FORCE TO PUSH LATCH HANDLE CLOSED. EXCESSIVE FORCE CAN DAMAGE LATCH. DO NOT ROTATE KEEPER EYE BOLT TO ADJUST LATCH TENSION. DAMAGE TO KEEPER MAY RESULT. If the force required to close the latch is not between 50 and 100 pounds, open the latch to relax the tension on the keeper. Adjust the force by rotating the adjustment star with a hex wrench or other suitable tool. The latch keeper mounting has an arrow to show the direction of rotation to increase the closing force. Properly adjusted latches close with a loud pop. Close the core cowl latches and check that all of the latch handles are flush with the core cowl panel contour. For Training Purposes Only

CAUTION:

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Figure 19 SCL

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Core Cowl Panel Latch Adjustment Page: 41

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

TURBINE EXHAUST SLEEVE AND PLUG General The turbine exhaust system makes a smooth exit path for turbine exhaust. The sleeve and plug form a nozzle to produce thrust The turbine exhaust sleeve is located aft of the turbine rear frame. The turbine exhaust plug is mounted inside the exhaust sleeve. Turbine Exhaust Sleeve The sleeve is conical, weighs 159 pounds (72 Kg) and is bolted to the turbine rear frame. It is acoustically treated with brazed titanium honeycomb. The core cowl rests on pads mounted around the sleeves leading edge. WARNING:

BE SURE FULL WEIGHT OF SLEEVE IS SUPPORTED BY CRADLE BEFORE REMOVING BOLTS. SLEEVE MAY SHIFT OR FALL INJURING PERSONNEL OR DAMAGING COMPONENTS.

Turbine Exhaust Plug The plug is also bolted to the turbine rear frame, weighs 33 pounds (15 Kg), and is a one piece construction. It is acoustically treated with brazed titanium honeycomb. BE SURE FULL WEIGHT OF PLUG IS SUPPORTED BEFORE REMOVING NUTS FROM UPPER HALF. PLUG MAY SHIFT OR FALL INJURING PERSONNEL OR DAMAGING COMPONENTS.

For Training Purposes Only

WARNING:

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ENGINE GENERAL

Figure 20 SCL

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Turbine Exhaust Sleeve and Plug Page: 43

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

FORWARD ENGINE MOUNT General The forward engine mount transmits thrust, vertical and lateral loads to the strut. Major component are the upper forward engine mount and the lower forward engine mount. Upper Forward Engine Mount The upper forward engine mount is part of the strut. It has holes for the tension bolts that attach it to the lower engine mount, and for the forward shear pin. Lower Forward Engine Mount The lower forward engine mount is made of a titanium alloy. The engine mount is attached to the aft inner flange of the fan frame. The mount has a platform which attaches to the fan frame with a failsafe clevis, a yoke bolted to the forward end of the platfom, two platfom links which attach to the yoke, and two frame links which attach the yoke to the fan frame. The ends of the yoke are also attached directly to the fan frame on both sides. One side has a tangential link to allow for thermal effects.

For Training Purposes Only

Engine Attachment The upper mount is attached to the lower mount platform with four tension bolts. Loads are transmitted from the fan case to the platform by the four links through the yoke. The tension bolts transmit vertical loads (the weight of the engine). A shear pin on the platform fits into the upper mount to transmit lateral loads (thrust) from the platform to the strut.

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ENGINE GENERAL

Figure 21 SCL

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Forward Engine Mount Page: 45

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

AFT ENGINE MOUNT General The aft engine mount transmits lateral, vertical and torque loads. Major components are the upper aft engine mount and the lower aft engine mount. Upper Aft Engine Mount The upper aft engine mount is part of the strut assembly. Two tamdem barrel nut assemblies in the mount connect to the tension bolts holding the upper and lower mounts together. Lower Aft Engine Mount The lower aft engine mount is attached to the engine turbine rear frame at two points. The left attachment has a tangential link between the mount and frame to allow for thermal effects. The mount is made of Titanium.

For Training Purposes Only

Engine Attachment The upper and lower mounts are connected together during engine installation using four tension bolts and barrel nuts. Two shear pins transmit lateral loads between the mounts.

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ENGINE GENERAL

Figure 22 SCL

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Aft Engine Mount Page: 47

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ENGINE GENERAL

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ENGINE VENTS AND DRAINS Drain Module A drain module collects leaked fluids and routes them to the drain mast. The module is mounted on the engine accessory gearbox. Access is through the thrust reverser halves. The accessories have separate drain inputs to the drain module. The drain module separates the leaked fuel from the leaked oil and hydraulic fluids. These leaked fluids are discharged during flight separately through the drain mast. When 200 Knots air speed is reached, a spring loaded valve to close ( no showed ) is opened by ram air from air inlet on the drain mast. This air flow empties the drain cavities by discharging accumulates fluids overboard through the mast There are push--to--open drain valves on the bottom of the module to help locate leakage sources. They are labeled to identify the different accessory seal drains. There are separate valves for the hydraulic pump pad, fuel / oil heat exchanger, main fuel pump pad, hydromechanical unit pad, starter pad and IDG pad.

For Training Purposes Only

Drain Mast The drain mast is mounted below the fan stator case and extends below the fan cowl. The mast drains the drain module and other accessories that are connected directly. The drain lines that exit directly through the drain mast are the strut drain, variable bypass valve actuators, variable stator vane actuators, fuel drain manifold, forward electrical junction box, IDG pressure relief valve, turbine air cooling valve actuators, fuel line shroud, and IDG over--temperature case drain. Scupper and Combustor Drains An oil tank scupper drain prevents service overflow (spillage) from accumulating on the engine. It is not connected to the drain mast. A combustor drain lets fluids drain from the combustor section when the engine is not running. It is not connected to the drain mast but has a line routed to the bottom of the rear turbine frame.

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ENGINE GENERAL

Figure 23 SCL

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Engine Vents and Drain Page: 49

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

POWER PLANT HAZARD AREAS Personnel must avoid the engine inlet and the exhaust area to prevent injury. The velocity of the fan discharge air is high enough to cause serious injury. When in reverse thrust, the fan air is discharged forward while the exhaust gas is discharged aft. A blast fence is recommended if the engines are going to be run for trim and power adjustment in an area where sufficient space is not available for dissipation of the fan and exhaust blast. High temperatures exist several hundred feet from the exhaust nozzle. Near the engine, the exhaust temperature is high enough to damage asphalt. Therefore, concrete aprons are suggested for run--up areas. WARNING:

DURING ENGINE RUN AT IDLE POWER, THE HAZARD ZONE MUST BE KEPT CLEAR, EXCEPT THAT ENGINE SAFETY BARRIER MAY BE SECURED IN INLET HAZARD ZONE.

WARNING:

FORWARD IDLE THRUST EXHAUST HAZARD ZONE MUST ALSO BE KEPT CLEAR DURING REVERSE THRUST OPERATION.

WARNING:

IF SURFACE WIND IS REPORTED GREATER THAN 25 KNOTS, INCREASE DISTANCE OF INLET BOUNDARY BY 20 PERCENT. IF RAMP SURFACES ARE SLIPPERY, ADDITIONAL PRECAUTIONS SUCH AS CLEANING THE RAMP WILL BE NECESSARY TO PROVIDE PERSONNEL SAFETY.

WARNING:

GROUND PERSONNEL MUST STAND CLEAR OF THESE HAZARD ZONES AND MAINTAIN COMMUNICATION WITH FLIGHT COMPARTMENT WITH FLIGHT COMPARTMENT PERSONNEL DURING ENGINE RUNNING.

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ENGINE GENERAL

Figure 24 SCL

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Power Plant hazard Areas 1 Page: 51

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NOTES :

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ENGINE GENERAL

Figure 25 SCL

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Power Plant Hazard Areas 2 Page: 53

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ENGINE GENERAL

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

ENGINE ENTRY CORRIDOR During engine operation, access to the engine may be required for maintenance purposes. The entry corridors to approach an operating engine are between the danger areas created by the inlet and exhaust flow. WARNING:

ALL PERSONNEL MUST AVOID DANGER AREAS IN FRONT AND REAR OF POWER PLANT AND REMAIN OUTSIDE OF ENGINE SAFETY BARRIER, IF USED, DURING GROUND RUNNING OPERATIONS. THE ENGINE IS CAPABLE OF DEVELOPING ENOUGH SUCTION AT THE INLET TO PULL A PERSON UP TO OR PARTIALLY INTO THE DUCT WITH POSSIBLE FATAL RESULTS. THEREFORE, WHEN APPROACHING ANY TYPE OF JET ENGINE, PRECAUTIONS MUST BE TAKEN TO KEEP CLEAR OF THE INLET AIR STREAM. THE SUCTION NEAR THE INLET CAN ALSO PULL IN HATS, GLASSES, LOOSE CLOTHING AND WIPE RAGS FROM POCKETS. ANY LOOSE ARTICLES MUST BE MADE SECURE OR REMOVED BEFORE WORKING AROUND THE ENGINE.

WARNING:

ENTRY CORRIDOR MUST BE USED ONLY UNDER FOLLOWING CONDITIONS: ENGINE OPERATION MAY NOT EXCEED LOW (MIN.) IDLE THRUST WHILE PERSONNEL ARE IN ENTRY CORRIDOR. POSITIVE COMMUNICATION BETWEEN PERSONNEL IN FLIGHT COMPARTMENT AND PERSONNEL USING ENTRY CORRIDOR IS MANDATORY.

WARNING:

INLET AND EXHAUST HAZARD AREAS MUST BE STRICTLY OBSERVED BY PERSONNEL IN ENTRY CORRIDOR.

WARNING:

IF SURFACE WIND IS REPORTED GREATER THAN 25 KNOTS, INCREASE DISTANCE OF INLET BOUNDARY BY 20 PERCENT. IF RAMP SURFACES ARE SLIPPERY, ADDITIONAL PRECAUTIONS SUCH AS CLEANING THE RAMP WILL BE NECESSARY TO PROVIDE PERSONNEL SAFETY.

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ENGINE GENERAL

Figure 26 SCL

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Engine Entry Corridor Page: 55

BOEING -- 767 / 300 CF6 -- 80C2 71 -- 00

ENGINE NOISE HAZARD AREA Jet engines produce noise capable of causing both temporary and permanent, loss of hearing. Even short exposures to extreme noise may result in damage to the ears. Noise affects the ear to cause unsteadiness or inability to walk or stand. EVEN WITH EAR PROTECTION, PROLONGED EXPOSURE CAN CAUSE EAR DAMAGE. All personnel must use ear protection. The cup--type ear protection is recommended. A chart for single--engine operation shows the limits of distance versus exposure time during different engine thrust conditions. WARNING:

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ENGINE GENERAL

Figure 27 SCL

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Engine Noise Hazard Areas Page: 57

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NOTES :

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ATA -- 72

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POWER PLANT

TABLE OF CONTENT General Airflow Stations Compressor Section Fan Module Fan Rotor Fan Booster Stator Case and Frame High Pressure Compressor Compressor Rear Frame and Combustor Turbine Modules Accessory Drives Module Accessory Gearbox Engine Borescope Inspection Ports

002 004 006 008 010 012 014 016 018 020 022 024

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ENGINE POWER PLANT

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POWER PLANT Description The CF6--80C2F is a two--shaft, axial flow, high bypass ratio turbofan engine. The fan and low pressure compressor (LPC) (five stages total), are shaft driven by a low pressure turbine (LPT). This assembly is called the N1 rotor, or fan rotor. The high pressure compressor (HPC) is shaft driven by a high pressure turbine (HPT) . This assembly is called the N2 rotor, or high pressure rotor. An annular combustor is used. The accessory gearbox is mounted to the core section of the engine.

For Training Purposes Only

Modules Five modules make up the engine. Each module may be replaced as an assembly without affecting engine performance or integrity. The five modules are: -- Fan module -- Core module -- High pressure turbine module -- Low pressure turbine module -- Accessory drives module

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ENGINE POWER PLANT

Figure 1 SCL

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Engine Summary Page: 3

BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

AIRFLOW STATIONS Specific positions along the engine air flow paths are assigned station numbers. These station numbers are used to identify pressure and temperature sensors . Air entering the core engine (LPC inlet) is called ”primary air flow“. Air flowing through the fan duct is called ”secondary air flow”. Major station designations include: -- Station 1.2: fan inlet at tip (secondary air flow) -- Station 1.4: fan outlet (secondary air flow) -- Station 2 : fan inlet at hub (primary air flow) -- Station 2.5: HPC inlet -- Station 3 : HPC outlet -- Station 4.9: LPT inlet -- Station 5 : LPT outlet Pressure sensors are identified with a P and the Temperature sensors are identified with a T. The ’point’ is dropped for sensor designation. For example, the pressure sensor at station 2.5 is called “P25“. Pressure and temperature sensors include: -- T12 fan inlet temperature -- P14 fan discharge pressure (secondary flow) -- P25,T25 HPC inlet pressure and temperature -- P3,T3 HPC discharge pressure (CDP) and temperature -- P49,T49 LPT inlet pressure, Temperature (EGT) -- T5 LPT discharge temperature.

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ENGINE POWER PLANT

Figure 2 SCL

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Airflow Stations Page: 5

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COMPRESSOR SECTION General The compressor section includes the fan module, and the high pressure compressor (HPC) section of the core module. Fan Rotor The fan rotor includes the fan rotor blades. The fan rotor blades serve two functions: -- First stage compression for air entering the LPC (primary flow) -- Acceleration of the air mass to develop about 80 % of the total engine thrust (secondary flow) -- Low Pressure Compressor (LPC) The LPC “boosts” (compresses) the air entering the HPC. There are five stages of compression: the fan rotor and the four stages of the LPC. The LPC is also called the booster. High Pressure Compressor (HPC) The HPC supplies high pressure air for combustion, engine cooling, and aircraft pneumatics requirements. The HPC is a 14--stage compressor. The first six stages have variable stator vanes (VSV).

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ENGINE POWER PLANT

Figure 3 SCL

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Compressor Section Page: 7

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ENGINE POWER PLANT

BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

FAN MODULE General The fan module is composed of the fan rotor, fan booster stator, fan case, and fan frame. Fan Rotor The fan rotor includes the fan rotor spinner, fan disk, 38 fan rotor blades, booster blades, a fan forward shaft and a fan mid shaft. The fan rotor spinner is black anodized aluminum. It is not anti--iced. The fan disk supports the fan rotor blades, the fan rotor spinner and a booster spool. The fan rotor blades are made of Titanium, and form a 93 inch diameter fan when installed. The blades are installed in dovetail slots in the fan disk. The booster blades make up stages 2 through 5 of the LPC. They are made of Titanium. The booster blades attach to the booster spool. The fan forward shaft supports and rotates the fan disk. Number 1 ball bearing and number 2 roller bearing support the fan shaft. The fan mid shaft is a tubular steel structure. It transmits torque from the LPT to the LPC. It is spline--coupled to the fan forward shaft and the LPT shaft.

Fan Frame The fan frame is the main structural component of the engine. The forward engine mount, the fan booster stator, and the aft fan case are attached to the fan frame. A number of other components are mounted on or supported by the fan frame struts.

For Training Purposes Only

Fan Booster Stator The fan booster stator is part of the LPC. The stator case directs the primary air flow into the HPC. Fan Case The fan case includes a forward fan case, and aft fan case. Kevlar cloth is wrapped around the forward fan case for fan blade containment. The cloth and an aluminum honeycomb core stiffen the fan case to prevent blade rubbing. The secondary flow (fan air) outlet guide vanes are part of the aft fan case.

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Figure 4 SCL

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Fan Module Page: 9

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FAN ROTOR Trim Balance Trim balance is required when the number 1 bearing or the turbine rear frame vibration level is out of limits. Both the spinner and the blade retainers have balance weights. Fan Rotor Spinner The fan rotor spinner is mounted to the fan disk with 38 bolts. One bolt hole is offset for indexing. Trim balance weights or screws are installed in 38 locations to help balance the engine. If the spinner is replaced, these balance weights and screws must be installed in the same location as on the old spinner. A seal ring helps stop air leaks. Fan Rotor Blades Blade 1 is installed in the second dovetail slot counterclockwise from the offset spinner bolt hole. The 38 blades are numbered counterclockwise from blade 1. The blades are installed by sliding them into the dovetail slots. A spacer, key and retainer hold the blade in the slot. A weight is added to the retainer for initial (coarse) balancing of the fan rotor. The ”moment weight class’ of the blade is stamped on the blade mounting platform. Blades of plus or minus one class are interchangeable without doing a fan trim balance. Opposite blades should be plus or minus one moment weight class. CAUTION:

ALL PARTS REMOVED, EXCEPT BOLTS AND NUTS, SHOULD BE MATCH MARKED OR NUMBERED FOR ASSEMBLY IN ORIGINAL ALIGNMENT AND POSITION. USE ONLY APPROVED MARKING MATERIAL.

ALL FIRST STAGE FAN BLADES, RETAINERS / SPACERS MUST BE INSTALLED BEFORE MEASURING BLADE TIP TO SHROUD CLEARANCES. When removing a fan blade it is necessary to remove the blade retainer, spacer and key from adjacent blades to allow enough blade movement to disengage the mid--span shroud.

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CAUTION:

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Refference Mark

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ENGINE POWER PLANT

Figure 5 SCL

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Fan Rotor Page: 11

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ENGINE POWER PLANT

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FAN BOOSTER STATOR, CASE AND FRAME Fan Booster Stator The fan booster stator case splits the air flow into primary and secondary paths. It includes fixed stator vanes for stages 2 through 4 of the LPC. The assembly bolts to the fan frame. Fan Case The fan case forms the outer shell of the fan duct (secondary airflow path) The forward fan case has an abradable shroud and Kevlar containment ring. The abradable shroud prevents fan blade / case damage if rubbing occurs. The Kevlar can hold a failed blade inside the engine. The 67 layers of Kevlar cloth are sealed and protected by a Kevlar / epoxy shell. The forward mating flange holds the fan inlet cowl. The aft fan case is bolted between the forward fan case and the fan frame. It forms the exit area for the secondary flow and has the stator vanes for LPC stage 5. The outlet guide vanes (OGVs) are aft of the fan blades, in the secondary flow. The OGVs are rigid graphite epoxy. Fan Frame The fan frame is a cast titanium hub with 12 radial struts welded to it. Strut number 1 is at the 12:00 position. The struts are numbered clock wise (i.e. strut 4 is at the 3:00 position).

For Training Purposes Only

Acoustic Liner Segment The acoustic liner segments reduce the noise level of the fan air exhaust. There are three bands of these segments in the inner wall. One band in the outer wall is forward of the fan blades in the forward fan case. The other bands are aft of the fan blades in the fan frame and case.

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ENGINE POWER PLANT

Figure 6 SCL

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Fan Booster Stator - Case and Frame Page: 13

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HIGH PRESSURE COMPRESSOR The high pressure compressor is part of the core module of the engine. The core module also includes the compressor rear frame assembly. The HPC case bolts to the fan module and to the compressor rear frame. The HPC is shaft--driven by the HPT. The shaft is supported by bearings 3R, 4R and 4B. The forward end of this shaft has a bevel gear to drive the accessory gearbox. The inlet guide vanes and the first five stages of the stator vanes are variable in angle. They are called variable stator vanes (VSV). Ports on the HPC case allow extraction of 7th, 8th, and 11th stage air for engine and airplane use.

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ENGINE POWER PLANT

Figure 7 SCL

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High Press Compressor Page: 15

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COMPRESSOR REAR FRAME / COMBUSTOR General The compressor rear frame (CRF) is the aft section of the core module. It houses the HPC outlet guide vanes, the combustor, and the HPT inlet guide vanes. Fuel nozzles and igniter plugs (not shown) are mounted in the CRF.

For Training Purposes Only

Combustor The combustor is an annulus formed by an inner liner and an outer liner. The liners are made of nickel alloys which have good strength characteristics at high temperatures. They are coated with a thermal barrier material to protect the parent metal.

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ENGINE POWER PLANT

Figure 8 SCL

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Compressor Rear Frame / Combustor Page: 17

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TURBINE MODULES General The HPT and the LPT are separate modules. Both are driven by exhaust gases from the combustor. The HPT drives the HPC. The LPT drives the LPC and fan blades. The turbine rear frame is part of the LPT. High Pressure Turbine The HPT includes the 2 stages of the HPT rotor and the stage 2 HPT nozzles. The stage 1 HPT nozzles are in the compressor rear frame. The rotor and blade assembly is cooled by a continuous flow of compressor discharge air. The nozzles are cooled by 11th stage compressor air. Low Pressure Turbine The low pressure turbine includes the 5 stage LPT rotor and stator, and the turbine rear frame. The turbine rear frame supports the turbine casing and bearing 6R. Bearing 6R supports the rotor. The first stage nozzle is cooled with 11th stage HPC air.

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Turbine Rear Frame The turbine rear frame is a major structural support. The aft engine mount is attached to the rear frame.

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ENGINE POWER PLANT

Figure 9 SCL

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Turbine Modules Page: 19

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ACCESSORY DRIVES MODULE The accessory drives module includes the accessory gearbox and the accessory heat shield. The gearbox is driven by the N2 rotor using gearboxes and drive shafts. An inlet gearbox located inside the fan module is driven by a bevel gear on the forward end of the N2 rotor shaft. A radial drive shaft transmits torque from the inlet gearbox to a transfer gearbox mounted on the fan frame under the compressor case. A horizontal drive shaft transmits torque from the transfer gearbox to the accessory gearbox. This drive shaft is enclosed in a housing. The accessory gearbox is mounted to the bottom of the compressor case. Selected pads on the gearbox have gear shaft adapters to make installation of accessories easier and more flexible. The heat shield is between the compressor case and the accessory gearbox to protect the accessories from the heat generated by the engine.

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ENGINE POWER PLANT

Figure 10 SCL

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Accessory Drives Module Page: 21

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ACCESSORY GEARBOX The accessory gearbox is a one--piece cast aluminium housing containing the bearings, shafts, gears, and oil nozzles needed to drive the accessories. Gearbox Forward Side The horizontal drive shaft enters the gearbox from the forward side. The forward side of the gearbox also has a drive pad to allow manual rotation of the engine for borescope use, etc. An access cover must be removed to use this drive. The following accessories and components are located on the forward side of the gearbox: -- Hydromechanical unit (HMU) -- N2 speed sensor -- Lube and scavenge pump assembly -- Control alternator -- Hydraulic pump A spare pad for a second hydraulic pump is available (not used).

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Gearbox Aft Side The following accessories and components are located on the aft side of the gearbox: -- Integrated drive generator (IDG) -- Pneumatic starter -- Fuel pump

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ENGINE POWER PLANT

Figure 11 SCL

JGB

May -- 2001

Accessory Gearbox Page: 23

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ENGINE POWER PLANT

BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

ENGINE BORESCOPE INSPECTION PORTS Engine internal inspection is primarily done by means of a borescope. The engine has borescope inspection ports for each stage the high pressure compressor, high pressure and low pressure turbine inlets, and stages 2 and 4 of the low pressure turbine. Additional borescope ports are in the compressor rear frame for the inspection of the combustion liner and first stage turbine nozzle. The N2 rotor is turned for borescope inspections by connecting a hand or motor power tool to a drive on the right forward face of the accessory gearbox. A cover is removed to use this drive. DO NOT INTERCHANGE BORESCOPE PORT PLUGS.

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NOTE:

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BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

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ENGINE POWER PLANT

Figure 12 SCL

JGB

May -- 2001

Engine Borescope Inspection Ports Right Side Page: 25

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

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ENGINE POWER PLANT

Figure 13 SCL

JGB

May -- 2001

Engine Borescope Inspection Ports Left Side Page: 26

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BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

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ENGINE POWER PLANT

Figure 14 SCL

JGB

May -- 2001

Power Plant Summary Page: 27

BOEING -- 767 / 300 CF6 -- 80C2 72 -- 00

NOTES :

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ENGINE POWER PLANT

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Page: 28

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

ATA -- 79 OIL SYSTEM TABLE OF CONTENT General Oil System Oil Storage Oil Servicing Oil Distribution Lub and Scavange Pump Supply and Scavange Inlet Screen Master Magnetic Chip Detector Servo Fuel Heater Fuel / Oil Heater Exchanger Scavange Oil Filter Distribution System Operation Oil Indicating System Oil Quantity Indicating Oil Pressure Indicating Oil Temperature Indicating Scavange Oil System Bypass Indicating Engine Oil System Schematic

002 004 006 008 010 012 014 016 018 020 022 024 026 028 030 032 034

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ENGINE OIL SYSTEM

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

ENGINE OIL SYSTEM General The oil system lubricates, cleans, and cools engine bearings and components. It has three major subsystems: -- storage -- distribution -- and indication. The oil system is separate from other engine and airplane fluid systems. Oil pressure is not regulated. Sensors and switches send signals to EICAS and the EEC indicating oil pressure, temperature, quantity, low oil pressure, and impending bypass of the scavenge oil filter.

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Component Location All oil system components are mounted on the engine. The oil tank is on the right side of the fan case. The scavenge oil filter is below the oil tank. Access to the oil tank and scavenge oil filter is through the right cowl panel. The lube and scavenge pump assembly is on the forward side of the accessory gearbox. An fuel / oil heat exchanger and servo fuel heater are on the lower right side of the engine near the accessory gearbox. Access to these *BREAK* oil system components is through the thrust reverser halves. Operation Oil flows by gravity from the oil tank to the lube and scavenge pump assembly. The accessory gearbox drives the lube and scavenge pump. The lube pump supplies oil to the engine bearings and gearboxes. Five scavenge pumps in the lube and scavenge pump assembly return the scavenge oil to the tank. The scavenge oil flows through the servo fuel heater, the fuel/oil heat exchanger, the scavenge oil filter and back into the oil tank.

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Page: 2

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 1 SCL

JGB

May -- 2001

Engine Oil System Page: 3

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

OIL STORAGE SYSTEM Oil Tank The oil tank stores the engine oil. The tank is aluminum with an external coating of silicone rubber for insulation. The tank volume is about 8 us. gallons (30.5 liters). When the system is properly serviced, the tank contains 6.6 U.S. gallons (25 liters) of oil. The tank includes pressure fill port connections, a sight glass and a drain plug. The interior has a deaerator surface (not show) to help remove air from the returning oil. Oil Tank Filler Cap The oil tank filler cap seals the manual fill port. The cap is on the upper right side of the oil tank. Access is either through the oil tank access door in the right fan cowl panel or by opening the panel. Oil Tank Pressurizing Valve The oil tank pressurizing valve maintains tank internal pressure. It is on top of the oil tank. The air--oil stream returning through the scavenge return tube pressurizes the oil tank. The valve keeps tank pressure at 7 to 11 psi above the transfer gearbox vent pressure.

For Training Purposes Only

Pressure Relief Valve The pressure relief valve is a back--up safety valve to relieve tank pressure. At 27 psi, it opens to ambient to prevent tank rupture. The valve is below the filler cap scupper basin.

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May -- 2001

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 2 SCL

JGB

May -- 2001

Oil Storage System Page: 5

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ENGINE OIL SYSTEM OIL SERVICING Oil Tank Maintenance Practices The oil level is normally checked between 5 minutes and 30 minutes after the engine has been shut down. For safety, do not check the oil for at least five minutes after shutdown. When filling manually, the tank is full when oil spills into the scupper basin. When pressure filling, the tank is full when oil flows through the overfill line. A sight glass is installed below the fill port scupper at about 3 quarts below full. The sight glass is not a reliable indication that the tank is properly serviced. The tank needs to be serviced if the ball in the sight glass is not at the top of the glass. If the ball is at the top of the sight glass, the tank is between two quarts low and full (or overfill). When an engine is motored, the scavenge pumps do not develop enough pressure to return oil to the tank. This causes oil to hide in the sumps and causes the sight glass to indicate that the oil level is low. Refer to the maintenance manual before servicing to prevent overfilling. When servicing the oil tank, check for the odor of fuel at the fill port. If there is fuel in the oil, replace the fuel / oil heat exchanger and the servo fuel heater, then drain and flush the engine oil system. After engine shutdown, the oil tank pressure slowly bleeds to ambient. WARNING:

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WAIT A MÍNIMUM OF FIVE MINUTES AFTER ENGINE IS SHUTDOWN BEFORE REMOVING FILLER CAP TO ALLOW TANK PRESSURE TO BLEED 0FF. HOT OIL GUSHING FROM THE TANK COULD CAUSE SEVERE BURNS.

Oil Tank Filler Cap Troubleshooting If the oil pressure indication changes with flight altitude, the O‘ring seal on the oil tank filler cap may be damaged. A bad seal causes low tank pressure that varies with barometric pressure. Oil pressure indication then changes with flight altitude due to varying atmospheric pressures.

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 3 SCL

JGB

May -- 2001

Oil Servicing Page: 7

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

OIL DISTRIBUTION SYSTEM General The oil distribution system supplies oil for lubricating the engine bearings and gearboxes. The oil is pressurized, cooled and filtered by the distribution system. Component Location The lube and scavenge pump is mounted to the front side of the accessory gearbox. The supply and scavenge inlet screens, internal pump components which are not shown, are located on the lube and scavenge pump. The magnetic chip detector is mounted in an oil tube adjacent to the drain module. The fuel / oil heat exchanger is bolted onto the fuel pump and is located aft of the hydromechanical unit (HMU). The scavenge oil filter is located below the oil tank of the fan case.

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General Operation The oil distribution system operation is automatic.

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ENGINE OIL SYSTEM

Figure 4 SCL

JGB

May -- 2001

Oil Distribution System Page: 9

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

LUBE AND SCAVENGE PUMP The lube and scavenge pump pressurizes the oil to assure ample oil flow to the bearings and gearboxes. The pump is mounted on the forward side of the accessory gearbox. It is driven by a spline shaft. Access is through the thrust reverser halves. The lube and scavenge pump has one pressure pump element and five scavenge pump elements. There are two rows of vane type positive displacement pumps in the pump housing. Each row has three pumping elements. The pumping elements have different capacities, determined by the diameter and length of each.

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Oil pressure is not regulated. All pump inlet ports are on the top surface except the pump drive shaft spline supply, which connects to a port on the underside. There is little space between the top of the pump and the underside of the engine. For easy removal and replacement, the oil tubes have flanges with threaded inserts which are secured by bolts that go through the pump body from the underside. Three reusable metal backed gaskets seal the tubes to the pump.

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Page: 10

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 5 SCL

JGB

May -- 2001

Lube and Scavenge Pump Page: 11

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

SUPPLY AND SCAVENGE INLET SCREENS The supply and scavenge inlet screens are in the lube and scavenge pump housing. Access is through the thrust reverser halves. Each inlet port to the six pump elements has a cleanable mesh finger screen to catch coarse debris. The supply inlet screen is 610 microns and the scavenge pump inlet screens are 940 microns. Each inlet screen is removed from the underside of the pump by unscrewing a hex cap. An optional magnetic chip detector can be installed in each screen through a threaded hole in the screen end cap.

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ENGINE OIL SYSTEM

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May -- 2001

Page: 12

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ENGINE OIL SYSTEM

Figure 6 SCL

JGB

May -- 2001

Supply and Scavenge Inlet Screens Page: 13

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

MASTER MAGNETIC CHIP DETECTOR The master magnetic chip detector attracts metal particles in the scavenge oil. The master magnetic chip detector is in the scavenge discharge flow tubing next to the drain module. Access is through the integrated drive generator service door (located on the left thrust reverser half inner cowl) or by opening the left thrust reverser half. The master magnetic chip detector is a permanent magnet probe. It is a 3--pinned bayonet--style probe with a knurled knob for installation and removal. A check valve in the housing permits removal of the chip detector probe without draining the oil system. Do not operate the engine without a chip detector installed, as the check valve can leak under these conditions.

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ENGINE OIL SYSTEM

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 7 SCL

JGB

May -- 2001

Master Magnetic Chip Detector Page: 15

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

SERVO FUEL HEATER The servo fuel heater cools the oil and heats the fuel used for hydromechanical unit (HMU) servo operations. The heater is bolted to a bracket on the right side of the accessory gearbox. Access is through the right thrust reverser half. The servo fuel heater has a multi--tube core mounted in a cylindrical housing that has two inlet ports and two outlet ports. One set of ports lets servo fuel pass through the tubes of the heat exchanger core. The other set of ports lets hot scavenge oil enter the heater through a relief valve assembly and flow around the fuel heater tubes. The relief valve opens at 60 pisd if the oil passage is blocked. Baffles change the oil flow direction four times before exiting the heater.

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ENGINE OIL SYSTEM

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Page: 16

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ENGINE OIL SYSTEM

Figure 8 SCL

JGB

May -- 2001

Servo Fuel Heater Page: 17

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

FUEL / OIL HEAT EXCHANGER The fuel / oil heat exchanger cools the oil and heats the fuel. it is bolted to the fuel pump on the bottom right side of the engine. Access is through the right thrust reverser half. The fuel / oil heat exchanger, like the servo fuel heater, has a multi--tube core mounted in a cylindrical housing that has two inlet ports and two outlet ports. One set of ports lets fuel pass through the tubes, of the heat exchangers core. The other set of ports lets oil pass around the core tubes inside the housing. All fuel always flows through the heat exchanger. A pressure relief valve opens at about 85--100 pisd to let scavenge oil bypass the core tubes. This bypass normally occurs during engine start in cold weather.

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ENGINE OIL SYSTEM

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May -- 2001

Page: 18

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 9 SCL

JGB

May -- 2001

Fuel / Oil Heat Exchanger Page: 19

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

SCAVENGE OIL FILTER The scavenge oil filter is mounted on a bracket attached to the fan stator case just below the oil tank on the right side of the tan case. Access is through the right tan cowl panel. The filter has an inlet port from the fuel / oil heat exchanger and an outlet port to the oil tank. The ports are labeled IN and OUT. The scavenge oil filter has a reversible disposable element. A relief valve lets oil bypass the filter. The valve begins to open at about 40 pisd. At 60 pisd, it is fully open. To change the filter element, the oil scavenge filter bowl is unscrewed from the filter head. Knurled bands on the bowl make it easier to grip the bowl for installation and removal. There are lugs at the bottom of the bowl so that a screwdriver can be used to loosen the bowl until it can be removed by hand. There is a shutoff valve in the filter head. When the filter is removed, the valve closes to prevent oil leakage. A new filter is installed by threading the filter and bowl into the filter head by hand until the shoulder seats against the head. A new packing must be used. The bowl is secured by lockwire through cast holes on the outside of the bowl.

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ENGINE OIL SYSTEM

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May -- 2001

Page: 20

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 10 SCL

JGB

May -- 2001

Scavenge Oil Filter Page: 21

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

DISTRIBUTION SYSTEM OPERATION Pressure Oil Flow The pressure pump element of the lube and scavenge pump supplies the oil to lubricate and cool the engine bearings and gears. Oil flows from the pressure pump through & check valve to the bearings and gears. Scavenge Oil Flow Three sumps collect the scavenge oil. The sumps are called the A sump, B / C sump, and the D sump. There are five scavenge pumps. These pumps service the accessory and transfer gearboxes and the B, C, and D sumps. Oil from the A sump drains down the radial drive shaft housing into the transfer gearbox. The oil from the sumps and gearboxes returns to the lube and scavenge pump through inlet screens with optional chip detectors to the scavenge pumps . The pumps return the oil to the tank through a common outlet. The magnetic chip detector, servo fuel heater, fuel / oil heat exchanger, and scavenge oil filter are in--line between the scavenge pumps and the tank.

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Abnormal Oil Flow Conditions The lubrication system works only when the engine is running. Motoring and windmilling operations do not supply adequate sump seal pressurization or sufficient scavenge flows. Because of this, apparent increased oil consumption rates and abnormal oil hiding occur.

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ENGINE OIL SYSTEM

Figure 11 SCL

JGB

May -- 2001

Distribution System Operation Page: 23

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

OIL INDICATING SYSTEM General The oil indicating system includes the oil quantity, oil temperature, oil pressure, low oil pressure and oil filter bypass indicating systems. Oil indications appear on EICAS. The secondary engine display and the PERF / APU page show oil pressure, temperature, and quantity. EICAS alert messages include L (R) ENG OIL PRESS and L (R) OIL FILTER. A L (R) ENG OIL PRESS light for each engine is located below the standby engine indicator. Most sensor signals are received directly by EICAS. The oil temperature signal is received by the EEC, which then sends the signal to EICAS.

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Sensors The components in the oil indicating system include: -- Oil Quantity Transmitter -- Oil Filter Differential Pressure Switch -- Oil Pressure Transmitter -- Low Oil Pressure Switch -- Oil Temperature Sensor

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 12 SCL

JGB

May -- 2001

Oil Indicating System Page: 25

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

OIL QUANTITY INDICATING SYSTEM Oil quantity appears on the EICAS secondary engine display and on the PERF/APU page. The oil quantity transmitter is mounted on a boss on top of the oil tank. The transmitter has a network of resistors, magnetic reed switches, and a floating permanent magnet which slides in a sensing unit tube. The magnetic float follows the oil level in the tank. Magnetic reed switches near the magnet close, changing the network resistance. EICAS sends a 28 volt dc reference voltage to the network and uses the response voltage to calculate the network resistance. Based on the resistance, EICAS determines and shows the oil level. The indication accuracy is +/-- 1 U.S. quart. The EICAS display has a low oil quantity white band indicating that the oil quantity is below 4 U.S. quarts. The transmitter cannot be adjusted by line maintenance.

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ENGINE OIL SYSTEM

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May -- 2001

Page: 26

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 13 SCL

JGB

May -- 2001

Oil Quantity Indicating System Page: 27

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

OIL PRESSURE INDICATING SYSTEM General Two independent oil pressure sensors (the low oil pressure switch and the oil pressure transmitter) send redundant oil pressure signals. The switch and transmitter are mounted to a bracket on the lube and scavenge pump. Both sensors measure the differential pressure between the lube and scavenge pump output and the accessory gearbox vent. Oil Pressure switch The low oil pressure switch is a diaphragm controlled, snap--action switch. The switch opens at 15 pisd and closes at 10 pisd. When the switch closes, the L (R) ENG OIL PRESS light is comes on and the EICAS alert message L (R) ENG OIL PRESS appears. This message is a level B message for CAA certified airplanes and a level C message for FAA certified airplanes. Oil Pressure Transmitter The oil pressure transmitter has a diaphragm that responds to pressure differential changes. A 28 V ac reference signal goes to the transmitter and to EICAS. The transmitter sends a bias signal to EICAS. EICAS converts the bias signal to oil pressure. Oil pressure appears on the secondary engine display and on the PERF / APU page.

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Oil Pressure Limits The lower red line limit for oil pressure is 10 pisd. The yellow band upper limit changes between idle and full power as a linear function of N2. The yellow band upper limit is 13 pisd when the engine is at low idle (60% N2). At full power (110% N2), the yellow band upper limit is 34 pisd.

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 14 SCL

JGB

May -- 2001

Oil Pressure Indicating System Page: 29

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

OIL TEMPERATURE INDICATING SYSTEM General The oil temperature sensor sends a signal to the EEC. The EEC sends a digital signal to EICAS. Oil temperature is indicated on the EICAS secondary engine display and on the PERF / APU page. Oil Temperature Sensor The oil temperature (TEO) sensor contains two chromel--alumel type thermocouples. The sensor is located on the forward side of the accessory gearbox immediately inboard and below the control alternator. The sensor mounts on a T--fitting in the scavenge oil return path between the master chip detector and the lube and scavenge pump.

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Oil Temperature Limits The operational range of the TEO sensor input to the EEC is from 81 to 352 ° F (--63 to 178 ° C) . The red line limit is 347 ° F (175 ° C). The yellow band range is from 320 ° F (160 ° C) to the red line limit.

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BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

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ENGINE OIL SYSTEM

Figure 15 SCL

JGB

May -- 2001

Oil Temperature Indicating System Page: 31

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ENGINE OIL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

SCAVENGE OIL FILTER BYPASS INDICATING SYSTEM Oil Filter Bypass When the differential pressure across the scavenge oil filter increases above 40 pisd, the bypass valve in the filter starts to open. Indication of impending bypass is given by the oil filter differential pressure switch.

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Oil Filter Differential Pressure Switch The oil filter differential pressure switch is a diaphragm controlled snap action switch that closes when the differential pressure across the scavenge filter element is more than 33 pisd. The switch is mounted to a bracket on the fan stator case below the oil tank and above the scavenge oil filter. An EICAS level C message L (R) OIL FILTER appears when the switch is closed.

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ENGINE OIL SYSTEM

Figure 16 SCL

JGB

May -- 2001

Scavenge Oil Filter Bypass Indicating System Page: 33

BOEING -- 767 / 300 CF6 -- 80C2 79 -- 00

ENGINE OIL SYSTEM SCHEMATIC Oil flows by gravity from the tank to the lube and scavenge pump assembly. This assembly has one pressure pump and five scavenge pump elements. The pressure pump element sends the oil under pressure to engine and gearbox bearings and gears. A check valve prevents reverse flow when the pump is not operating. The scavenge pump outflows are combined. The oil goes past the oil temperature sensor and the magnetic chip detector, through the servo fuel heater and the fuel / oil heat exchanger, and then through the scavenge oil filter. The scavenge oil filter removes contaminants from the oil. Oil returns to the tank through a deaerator. A pressurizing valve and a pressure relief valve maintain proper oil tank pressure. Sensors, transmitters and switches supply indications of oil quantity, oil temperature, oil pressure, low oil pressure and an impending bypass of the scavenge oil filter.

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ENGINE OIL SYSTEM

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Page: 34

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ENGINE OIL SYSTEM

Figure 17 SCL

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May -- 2001

Engine Oil System Schematic Page: 35

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NOTES :

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ENGINE OIL SYSTEM

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Page: 36

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

ATA -- 73 FUEL SYSTEM TABLE OF CONTENT General Fuel System Component Locations Fuel Distribution -- Schematic Main Fuel supply Hose Fuel Pump Fuel Pump Operation Fuel Filter and Element Servo Fuel Heater Fuel Manifold and Tubes Fuel Nozzles Combustor Drain Valve Fuel Indicating System -- Schematic Fuel Flow Transmitter Fuel Pump Interstage Transmitter Fuel Filter Differential Pressure Switch Engine Fuel System Operation

002 004 006 008 010 012 014 016 018 020 022 024 026 028 030 032

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ENGINE FUEL SYSTEM

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Page: 1

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

ENGINE FUEL SYSTEM General The engine fuel system includes distribution, control and indication. The control functions are covered in the engine control chapter. Distribution The fuel distribution system receives and pressurizes fuel from the airplane fuel tanks. The fuel is heated by engine oil in the fuel / oil heat exchanger and then filtered. After being filtered, the fuel is heated in the IDG fuel / oil heat exchanger and distributed through the fuel tubes to the fuel nozzles in the engine combustor. A servo fuel heater provides additional heat to the servo fuel used by the hydromechanical unit (HMU) for control. Control The hydromechanical unit (HMU) provides fuel metering and engine air systems control functions. Operation of the HMU is covered in the engine control chapter.

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Indication Fuel flow rate is displayed on EICAS using a fuel flow transmitter. Fuel pump interstage pressure is displayed on EICAS, using a fuel pump interstage pressure transmitter. Impending blockage of the fuel filter is indicated by the EICAS status message L (R) ENG FUEL FILT, using a fuel filter differential pressure switch.

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ENGINE FUEL SYSTEM

Figure 1 SCL

JGB

May -- 2001

Engine Fuel System Page: 3

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

ENGINE SYSTEM COMPONENT LOCATIONS The fuel distribution system pressurizes, filters, and distributes fuel. It delivers fuel from the airplane fuel tanks to the engine combustion section. It also supplies pressurized and heated fuel for use by the engine air system. Component locations are as follows: -- Main fuel supply hose: routed from strut down right side of engine to fuel pump inlet port. -- Fuel pump: mounted on right aft side of accessory gear box. -- Fuel oil heat exchanger: mounted on bottom side of fuel pump. -- Fuel filter: mounted on outboard side of fuel pump. -- Servo fuel heater: mounted to heatshield on right side above accessory gearbox. -- IDG fuel / oil heat exchanger: supported by brackets attached to the right underside of the accessory gearbox. -- Fuel tubes (manifold): mounted around the combustor connecting to the fuel nozzles. -- Fuel nozzles: installed evenly around the combustor. -- Fuel filter differential pressure switch: bracket mounted to top of the fuel filter. -- Fuel pump interstage pressure transmitter: mounted in fuel pump port. -- Fuel flow transmitter: supported by brackets attached to right outer corner of the accessory gearbox.

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ENGINE FUEL SYSTEM

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Page: 4

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ENGINE FUEL SYSTEM

Figure 2 SCL

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May -- 2001

Engine Fuel System - Component Location Page: 5

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL DISTRIBUTION SYSTEM - SCHEMATIC Interfaces Two systems interface with the fuel distribution system: -- engine oil -- IDG oil. Engine oil is cooled by fuel in the fuel / oil heat exchanger. Engine oil is also used to heat the engine fuel system servo fuel in the servo fuel heater. IDG oil is cooled by fuel in the IDG fuel/oil heat exchanger.

For Training Purposes Only

General Operation Fuel from the airplane fuel system flows through the main fuel supply hose into the fuel pump. The pump pressurizes the fuel and discharges it through the fuel/oil heat exchanger and the fuel filter to the HMU. The fuel, metered by the HMU, flows through the fuel flow transmitter, the IDG fuel / oil heat exchanger and the fuel tubes to the fuel nozzles. The fuel nozzles spray the fuel into the combustion chamber for combustion. A portion of the fuel filter outflow is directed through the servo fuel heater to the HMU for internal use and control of the engine air system.

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ENGINE FUEL SYSTEM

Figure 3 SCL

JGB

May -- 2001

Fuel Distribution System Page: 7

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

MAIN FUEL SUPPLY HOSE The main fuel supply hose connects the airplane fuel supply line (in the engine strut) to the fuel pump (on the engine) It is on the right side of the engine core section. Access is through the right thrust reverser half. The hose is connected at the strut with a coupler. A mounting flange connects the hose to the fuel pump. Four clamps hold the hose to the engine between the strut and pump. An insulation blanket surrounds part of the hose to protect the system from thermal effects. The fuel supply hose is drained prior to disconnect by two drain plugs on the fuel pump. CATCH THE DRAINED FUEL USING A SUITABLE 5 GALLON CAPACITY CONTAINER.

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NOTE:

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ENGINE FUEL SYSTEM

Figure 4 SCL

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL PUMP General The fuel pump pressurizes the fuel. The fuel pump attaches to the aft right pad of the accessory gearbox using an adapter with a hinged V flange coupling. A spline drive shaft engages the pump to the accessory gearbox adapter using an O‘ring seal. A carbon seal (not shown) keeps fuel out of the accessory gearbox. A cleanable metal interstage strainer protects the pump from particle damage. A fuel discharge port and a fuel return port connect the fuel pump to the HMU. Two drain plugs are located on the bottom of the pump. The fuel / oil heat exchanger, fuel pressure transmitter, and fuel filter are mounted to the pump assembly. Two ports on the pump connect a fuel filter differential pressure switch.

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Removal and Installation The pump may be removed with the heat exchanger and filter attached if desired. To avoid damage to the seals, do not allow the pump assembly to hang from the drive shaft during removal or installation. The fuel pump weighs approximately 43 pounds.

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ENGINE FUEL SYSTEM

Figure 5 SCL

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ENGINE FUEL SYSTEM

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FUEL PUMP OPERATION Operation The fuel pump has both an impeller (interstage) pump and a gear pump. Both pumps are driven by a common spline drive from the accessory gearbox. The impeller pump pressurizes fuel to prevent gear pump cavitation. The gear pump generates the high pressure and flow to supply the HMU and fuel nozzles. The fuel flows from the impeller pump through the interstage strainer to the positive--displacement gear pump. The impeller pump discharge (boost) pressure is 0--152 psid, depending on RPM. The gear pump discharge (outflow) pressure is maintained below 1500--1700 psid by a relief valve. Fuel flows from the gear pump through the heat exchanger and fuel filter to the discharge port (to the HMU). Excess fuel from the HMU enters the pump through the return port (located between the impeller and gear pump stages). The fuel pump interstage pressure transmitter measures the interstage fuel pressure for indication on EICAS.

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Servicing The metal interstage strainer is removed for cleaning. If the strainer is clogged, N2 generally does not increase above 45--50%.

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ENGINE FUEL SYSTEM

Figure 6 SCL

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Fuel Pump Operation Page: 13

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL FILTER AND ELEMENT General The fuel filter removes particles large enough to cause contamination or damage. The filter assembly is bolted to th,e outboard side of the fuel pump. A servo fuel outlet port is on the filter housing. The servo fuel flows through a wash screen with a relief valve. The valve opens at about 15 psid if the screen becomes blocked. The filter element is a disposable 10 Micron nominal (35 micron absolute) unit. A coarse aluminum mesh supports a pleated epoxy impregnated glass / polyester compound. Each end has a seal ring. A relief valve lets fuel bypass a clogged filter element at about 35 psid. Removal and Installation To replace the filter element, unscrew the filter bowl from the housing. Either end of the fuel filter element can be put into the filter bowl. Install the filter bowl and tighten by hand. DO NOT OVERTIGHTEN FUEL FILTER BOWL.

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NOTE:

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ENGINE FUEL SYSTEM

Figure 7 SCL

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Fuel Filter and Element Page: 15

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

SERVO FUEL HEATER The servo fuel heater heats the fuel used for HMU servo operations to prevent icing of the fuel. The heater is bolted to a bracket in the accessory compartment on the right side of the accessory gearbox. Hot oil from the engine lube system enters the heater through a relief valve assembly to flow around the fuel heater tubes. The relief valve opens at 60 psid if the oil passage becomes blocked. Baffles force the oil to change flow direction four times before exiting the heater. Fuel passes straight through the heater tubes without bypass, absorbing heat from the oil before exiting. Included in the assembly is a second valve at the heat exchanger fuel outlet to limit heat input to the fuel. If the fuel becomes to hot (88 -- 93 ° C) the thermal unit closes off the servo oil return to the gearbox. This causes a differential pressure across the oil bypass valve. The bypass valve moves and allows inlet oil to proceed directly to the outlet oil port. The exchange of heat from the oil to the fuel is thereby stopped.

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ENGINE FUEL SYSTEM

Figure 8 SCL

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Servo Fuel Heater Page: 17

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL MANIFOLD AND TUBES General A fuel manifold and tubes carry metered fuel from the HMU and IDG fuel / oil heat exchanger to the 30 fuel nozzles. The manifold encircles the engine at the combustion section. The fuel manifold is divided into two segments. Each segment carries fuel to 15 fuel nozzles using individual fuel supply tubes welded to the manifold. The manifold segments are connected with couplings at the 6:30 and 12:30 positions. The tube--to--fuel nozzle couplings are covered by a shroud to catch leakage. The shrouds are connected to a drain manifold. The fuel collected in the drain manifold is routed to the drain mast where it is discharged.

For Training Purposes Only

Removal and Installation Loosen the knurled nuts at the nozzle and at the drain manifold. Slide the shroud aft. This exposes the shroud--to--nozzle packing and the connection between the fuel nozzle and fuel tube.

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ENGINE FUEL SYSTEM

Figure 9 SCL

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Fuel Manifold and Tubes Page: 19

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL NOZZLES General The fuel nozzles distribute and atomize the fuel in the combustor. The 30 nozzles are mounted through the compressor rear frame (CRF) and are numbered 1 through 30 clockwise from the top. Access to the nozzle is through the thrust reverser halves. Nozzles 15 and 16 have larger primary flow passages that supply a richer flow to prevent flameout. They are identified by blue bands and have a different part number from the nozzle used at all other locations; the other 28 nozzles have aluminuin colored bands. Operation Fuel enters the nozzles through an inlet check valve. The valve opens at 20 psid and keeps the fuel manifold from draining into the combustor when the engine is shut down. At low fuel flows, a flow divider valve directs fuel to the primary flow passage. As fuel flow increases, the flow divider valve opens to let fuel enter the secondary flow passage.

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Removal and Installation Replace the fuel nozzle with one with the same color band and part number. Disconnect the fuel and drain manifolds before disconnecting nozzles. When replacing nozzles under the engine (nozzles 9 through 22), the metallic gasket can be taped to hold it in place during installation. Remove the tape before final tightening.

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ENGINE FUEL SYSTEM

Figure 10 SCL

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Fuel Nozzles Page: 21

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

COMBUSTOR DRAIN VALVE General The combustor drain valve lets fuel or other liquids drain from the combustor and CRF when the engine is shutdown. The valve is at about the 5:30 position, clamped to the LPT cooling air manifold. Access is through the right core cowl. The combustor drain valve is a spring--loaded to open poppet valve. A forward tube connects the valve to a fitting at the 6:00 position on the CRF. An aft tube carries drainage overboard near the exhaust sleeve.

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Operation When the engine is running, combustor gas pressure closes the drain valve. When the engine is shut down, the valve opens to drain fluids.

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ENGINE FUEL SYSTEM

Figure 11 SCL

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Combustor Drain Valve Page: 23

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL INDICATING SYSTEM SCHEMATIC The fuel indicating system supplies indications of the engine fuel system operation to the flight crew. The indications include fuel flow, fuel pump interstage pressure and fuel filter bypass warnings. All the engine fuel system sensors are mounted on the engines. The indications on the flight deck normally appear on the lower EICAS display. These include flows on the secondary engine parameter display, flows and pressures on the PERF / APU page, and messages on the status and ECS/MSG pages. In addition, fuel flow is sent to the flight management computer (FMC).

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ENGINE FUEL SYSTEM

Figure 12 SCL

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Fuel Indicating System - Schematic Page: 25

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL FLOW TRANSMITTER General The fuel flow transmitter measures the fuel mass flow rate to the fuel nozzles. The transmitter is near the right side of the accessory gearbox below the fuel pump. The input to the transmitter comes from the HMU. The output goes to the IDG fuel / oil heat exchanger. Access is through the right thrust reverser half. Operation The transmitter has a flow director, swirl generator, rotor, and turbine. The rotor spins freely and has two magnets that create pulses in start and stop coils. The turbine can turn but is kept from spinning by a restraining spring. Incoming fuel goes through the flow director and is given angular momentum by the swirl generator, making the rotor spin. One of the magnets on the rotor generates a signal in the start coil. The other magnet creates a signal in the stop coil by passing under the signal blade attached to the turbine. The amount of time between the start and stop signals varies in proportion to the fuel flow rate. The start and stop pulses are received by the EEC, which then calculates the fuel flow rate. The EEC sends digital flow rate information to EICAS.

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Unique Practices Transmitters removed from the airplane and not reinstalled within 24 hours must be protected against internal corrosion. Fill the transmitter with enough engine oil to coat all parts, drain the oil, and install protective covers (not shown) on both ends.

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ENGINE FUEL SYSTEM

Figure 13 SCL

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Fuel Flow Transmitter Page: 27

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL PUMP INTERSTAGE PRESSURE TRANSMITTER The fuel pump interstage pressure transmitter measures the interstage pressure in the fuel pump. It is mounted on the fuel pump next to the fuel filter. The transmitter is a variable reluctance unit. It sends an electrical analog signal to EICAS. EICAS calculates the fuel pressure and shows the pressure on the PERF / APU page.

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ENGINE FUEL SYSTEM

Figure 14 SCL

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Fuel Pump Interstage Pressure Transmitter Page: 29

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ENGINE FUEL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

FUEL FILTER DIFFERENTIAL PRESSURE SWITCH General The fuel filter differential pressure switch closes to indicate an excessive difference in fuel pressure across the fuel filter. The switch is mounted on the fuel pump. Access is through the right thrust reverser half.

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Operation The switch closes when the differential pressure across the filter is greater than 23 psid. A latched EICAS status and maintenance message L (R) ENG FUEL FILT appears after a 10 second time delay. If the differential pressure decreases to less than 19.5 psid within 10 seconds after the switch closes, the switch opens and the message goes away. The filter bypass valve does not open until about 35 psid. The EICAS message shows impending fuel filter bypass and does not necessarily indicate that the bypass valve is open.

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ENGINE FUEL SYSTEM

Figure 15 SCL

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Fuel Filter Differential Pressure Switch Page: 31

BOEING -- 767 / 300 CF6 -- 80C2 73 -- 00

ENGINE FUEL SYSTEM OPERATION Fuel enters the engine fuel system from the fuel tanks through the main fuel supply hose. The fuel pump is spline shaft driven by the accessory gearbox, and pressurizes the fuel using an impeller (boost) pump and a gear pump. The fuel pump interstage pressure transmitter measures impeller pump pressure. The fuel is heated in the fuel / oil heat exchanger and filtered by the fuel filter. A fuel filter differential pressure switch sends a signal to EICAS if the filter is becoming blocked. The HMU meters the fuel to the nozzles for combustion. The HMU gets a separate supply of servo fuel from a port on the fuel filter. The servo fuel is heated by the servo fuel heater. The fuel flow transmitter measures the fuel mass flow rate for EICAS indication. The IDG fuel / oil heat exchanger transfers additional heat to the combustion fuel. The fuel manifolds and tubes carry the fuel to the fuel nozzles. The tube and nozzle couplings have a shroud and a drain manifold to deliver leakage to the drain mast. The nozzles atomize the fuel for combustion. When the engine is shut down, the combustor drain valve opens to let fluids in the combustion section drain.

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ENGINE FUEL SYSTEM

Figure 16 SCL

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Engine Fuel System Operation Page: 33

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NOTES :

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ATA -- 75 AIR SYSTEM TABLE OF CONTENT Air System General Air System Component Locations Engine Cooling System CCCV System CCCV Control TCC System TCC Control Compressor Discharge Temp. Sensor Compressor Airflow Control System VSV System Components VBV System Components VSV and VBV Control Engine Air System Indications Engine Air System Operation

002 004 006 008 010 012 014 016 018 020 022 024 026 028

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ENGINE AIR SYSTEM

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ENGINE AIR SYSTEM

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ENGINE AIR SYSTEM General The engine air system controls air through the compressor, and controls airflow for engine and accessory cooling. The EEC and the HMU control these systems. Engine Cooling Systems Fan discharge air is used to cool the engine using two systems: -- Core Compartment Cooling System -- Turbine Case Cooling (Active Clearance Control) System

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Compressor Airflow Control Systems LPC discharge air entering the HPC is regulated by two systems: -- Variable Bypass Valves (VBV) -- Variable Stator Vanes (VSV) Air from the HPC is used to meet service bleed demands, to cool the ignitor leads, and to supply air to the aircraft and engine anti--ice systems. These air systems are covered elsewhere in the course material.

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ENGINE AIR SYSTEM

Figure 1 SCL

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Engine Air Systems Page: 3

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ENGINE AIR SYSTEM

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ENGINE AIR SYSTEM COMPONENT LOCATIONS Engine Cooling Systems A core compartment cooling valve (CCCV) mounted on the left side of the engine core controls fan air to a manifold used for accessory cooling. An CCCV solenoid, integral to the CCCV valve, controls the operation of the CCCV in response to EEC commands. A high pressure turbine cooling (HPTC) valve mounted on the right side of the diffuser case controls HPT blade tip clearance. The HPTC manifold encircles the HPT case.

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Compressor Airflow Control Systems Variable bypass valve (VBV) actuators mounted on each side of the fan frame control the position of the bypass valves. Variable stator vane (VSV) actuators mounted on each side of the forward HPC case control the positions of the HPC variable inlet guide vanes and stator vanes.

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ENGINE AIR SYSTEM

Figure 2 SCL

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Engine Air System Component Location Page: 5

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ENGINE AIR SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 75 -- 00

ENGINE COOLING SYSTEM General The engine cooling system supplies external cooling air to the engine and accessories. Valves control the cooling air flow to maximize engine efficiency. Core Compartment Cooling Valve (CCCV). The CCCV controls fan air used to cool engine accessories. The EEC controls the position of the valve through the CCCV solenoid. The valve is closed at high power and high altitudes. The valve is located on the left side of the engine.

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Turbine Case Cooling Valve A high pressure turbine cooling (HPTC) valve controls air flow through the HPTC manifold. The manifold blows fan air on the surface of the turbine case to control the case thermal growth. The valve is powered by HMU servo fuel, and controlled by the EEC through an electro hydraulic servo valve (EHSV) located inside of the HMU.

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ENGINE AIR SYSTEM

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Engine Cooling System Page: 7

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ENGINE AIR SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 75 -- 00

CCCV SYSTEM General The CCCV system supplies controlled cooling air for the core mounted engine accessories. The system conserves primary air by reducing the core cooling at low power and high altitudes. The system has one core compartment cooling valve (CCCV). The valve is controlled by a CCCV solenoid integral to the valve. The solenoid is controlled by the EEC. Core Compartment Cooling Valve (CCCV) The core compartment gets fan air for cooling through the CCCV and manifold. The valve is at the 10:00 position on the HPC case. The butterfly--type valve is spring--loaded to open. When the valve is open, airflow is not restricted. It closes when eleventh--stage air is sent to the diaphragm in the valve actuator. When the valve is closed, the cooling airflow is reduced, but not cut off completely. A position indicator on the actuator gives a visual indication of valve position. The manifold sends airflow to the HPC case, the IDG, hydraulic and fuel pumps, and other accessories.

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CCCV Solenoid The CCCV solenoid controls the flow of eleventh stage air that controls the CCCV. The solenoid valve is spring--loaded to closed. The eleventh stage air pressure comes from the ESCV supply duct on the left side of the engine. When the solenoid is energized, the eleventh stage air pressure goes to the CCCV to close it.

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ENGINE AIR SYSTEM

Figure 4 SCL

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CCCV System Page: 9

BOEING -- 767 / 300 CF6 -- 80C2 75 -- 00

CCCV CONTROL The EEC, through the CCCV solenoid, controls the flow of eleventh--stage air used to close the CCCV. The solenoid has two electrically independent coils, each commanded by a different channel of the EEC. The EEC energizes the CCCV solenoid to close the valve when the conditions below are met: -- N1 greater than 86 %. -- Ambient pressure less than 7.95 PSIA (approxi. 17,000 Ft. of altitude). -- T49 (EGT) less than 699 ° C. -- The engine acceleration rate is less than 70 RPM per second. -- The commanded N2 is not more than 5 % greater than the actual N2. The active EEC channel energizes the CCCV solenoid closing the core compartment cooling valve. There is no position feedback from the CCCV.

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ENGINE AIR SYSTEM

Figure 5 SCL

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CCCV Control Page: 11

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ENGINE AIR SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 75 -- 00

TURBINE CASE COOLING (ACTIVE CLEARANCE CONTROL) Description The turbine case cooling (active clearance control) system uses separate manifolds to cool the LPT and HPT cases. The fan air to the HPT manifold is controlled by the high pressure turbine cooling (HPTC) valve. The LPTC and HPTC manifolds encircle and direct fan air onto their respective turbine cases. This reduces cases expansion, thus minimizing turbine blade tip to case clearance which increases turbine efficiency. The HPTC valve is mounted on the right side of the engine at the 1:00 position near the eleventh--stage bleed manifold. The valve is clamped at each end to the respective cooling air pipes through which they receive fan air.

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HPTC Valve The HPTC valve is a butterfly--type valve controlled by a hydraulic piston actuator. Modulation of the valve is controlled by hydraulic fluid pressures received from electro--hydraulic servo valve (EHSV) in the hydromechanical unit (HMU). The EHSV is controlled by the EEC. The valve assembly has two linear variable differential transformers (LVDT‘s) which supply valve position signals to he EEC. There is an electrical connector for each LVDT. One LVDT is excited and read by EEC channel A. The other LVDT is excited and read by EEC channel B.

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ENGINE AIR SYSTEM

Figure 6 SCL

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Turbine Case Cooling (ACC) Page: 13

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TURBINE CASE COOLING CONTROL HPT cooling is controlled by the active channel processor within the EEC, the HPTC EHSV within the HMU, and the HPTC valve. The active clearance control software components inside the EEC channel processor are the dimensional calculators, command calculators, demand calculators, and valve drivers. The dimensional calculators issue a ”size error” signal whenever the calculators determine that the clearance between turbine case and turbine blade tip are incorrect. To do these calculations, the dimensional calculators use several temperature, pressure and speed parameters. The command calculators receive the ”size error” signals and convert them to valve position command signals. The valve position command signal is a percentage, with 0 % equal to valve--fully--closed and 100 % equal to valve--fully--open. Using the valve feedback signals, the demand calculator determines the error between the actual and commanded valve positions, and generates an output equal to the error. The error signal is sent to the valve driver which converts the digital signal to a DC signal. This signal goes to the EHSV in the HMU where it controls the position of the HPTC valve.

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ENGINE AIR SYSTEM

Figure 7 SCL

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Turbine Case Cooling Control Page: 15

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COMPRESSOR DISCHARGE TEMPERATURE (T3) SENSOR The T3 temperature sensor measures HPC discharge air temperature. The EEC uses this temperature to sequence the bore cooling valves (BCV‘s) and the active clearance control valves. The T3 temperature sensor is mounted to the forward end of the compressor rear frame at the 11:30 position. The T3 sensor has dual chromel / alumel thermocouples, one for each EEC channel. A single electrical connector sends both outputs to the cold junctions inside the EEC. The connector is located above the EGT shunt junction box on a bracket on the LPT cooling air tube. The outputs from the T3 sensor go to the connector through a metal cased ceramic--sheathed lead. The operational range of the T3 input to the EEC is from --75 to +1337 ° F (--60 to +725 ° C).

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ENGINE AIR SYSTEM

Figure 8 SCL

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Compressor Discharge Temperature (T3) Sensor Page: 17

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ENGINE AIR SYSTEM

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COMPRESSOR AIRFLOW CONTROL SYSTEM General The compressor control system prevents compressor stall (surges) and improves engine efficiency. Two systems, the variable stator vanes (VSV) and variable bypass valves (VBV), control the HPC airflow. Both systems use hydraulic actuators. Servo fuel from the HMU is used as the hydraulic fluid to control the actuators. The variable stator vanes include the HPC inlet guide vanes and the first five stages of the HPC stator vanes. Modulation of these vanes permits optimum compressor performance throughout the engine operating range. The VSV components are located on the forward HPC case. The VSV‘s are varied in unison by two VSV actuators. They are closed at low power and modulate open as power increases. The VBV components are in the fan frame. Twelve valves are modulated in unison by two actuators. The VBVs are open at low power and modulate toward closed as power increases. The open valves divert a portion of the LPC primary discharge from the HPC to the secondary flow path. Each VSV and VBV actuator has a linear variable differential transformer (LVDT) to send feedback signals to the EEC. The actuator LVDT‘s on the left side of the engine are excited by and send feedback signals to EEC channel A. The right side actuator LVDT‘s are excited by and send feedback signals to EEC channel B. Operation The EEC uses input signals from engine sensors to control electro--hydraulic servo valves (EHSV‘s) in the HMU. The EHSV‘s use servo fuel to modulate the VSV and VBV actuators. The EEC increases current to the EHSV in proportion to N2. The EHSV directs servo fuel pressure to the actuators to move them to the commanded position.

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ENGINE AIR SYSTEM

Figure 9 SCL

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Compressor Airflow Control System Page: 19

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VARIABLE STATOR VANE (VSV) SYSTEM COMPONENTS General The VSV system components include two actuators, two actuation levers, and six actuation rings connected to VSV lever arms. Access to the VSV system components is through the thrust reverser halves. VSV Actuators The VSV actuators are a double--action piston type, mounted at 3:00 and 9:00 positions on the HPC case forward flange.

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Operation The HMU sends high--pressure servo fuel to the head and rod ends of the VSV actuators. Increasing the head end servo fuel pressure, and decreasing the rod end fuel pressure, causes the actuator pistons to extend. This causes the left actuation lever to lower, the right actuation lever to raise, and the actuation rings to rotate counterclockwise (aft looking forward), opening the VSV‘s. Increasing the rod end servo fuel pressure and decreasing the head end servo fuel pressure closes the VSV‘s. An electrical connector on each actuator provides position feedback to the EEC from an LVDT located inside the actuator. The left actuator LVDT is excited by and sends position feedback signals to EEC channel A. The right actuator LVDT is excited by and sends position feedback signals to EEC channel B.

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ENGINE AIR SYSTEM

Figure 10 SCL

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Variable Stator Vane System Components Page: 21

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VARIABLE BYPASS VALVE (VBV) SYSTEM COMPONENTS General The VBV system components include two actuators, a unison ring, bell cranks, and 12 bypass valves. Access to the VBV system components is through the thrust reverser halves. VBV Actuators The VBV actuators are double--action piston type, mounted near the 4:00 and 10:00 positions on the fan frame. The 12 VBV‘s are equally spaced around the LPC case between the fan frame struts. They are rectangular metal plates that cover the bypass valve outlets. LPC primary discharge air is diverted through open VBV‘s into the secondary air flow path.

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Operation A unison ring interconnects all12 VBV‘s using bell cranks. All VBV‘s operate in unison in response to the actuators. The HMU sends high pressure servo fuel to the head and rod ends of the VBV actuators. Increasing the head end servo fuel pressure, and decreasing the rod end fuel pressure, causes the actuator pistons to extend. This causes the unison ring to rotate counterclockwise, opening the VBV‘s increasing the rod end servo fuel pressure and decreasing the head end servo fuel pressure closes the VBV‘s. An electrical connector on each actuator provides position feedback to the EEC from a LVDT located inside the actuator. The left actuator LVDT is excited by and sends position feedback signals to EEC channel A. The right actuator LVDT is excited by and sends position feedback signals to EEC channel B.

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ENGINE AIR SYSTEM

Figure 11 SCL

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Variable Bypass Valve System Components Page: 23

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VSV AND VBV CONTROL General The logic schedules for VSV and VBV control are incorporated into the EEC software. The VSV‘s are modulated as a function of actual N2, T25 and PO. The VBV‘s are modulated as a function of actual N1, TAT and the VSV positions. When the engine is started, the VBV‘s are open and the VSVs are closed. As the engine accelerates, the EEC commands the EHSV to signal the VSV actuators to gradually open the vanes. The position feedback signal tells the EEC that the actuators have moved to the commanded position. The VSV position is also used by the EEC to schedule the position of the VBV‘s. The VBV actuators get fuel pressure signals to gradually close as power increases. At high power, the VSV‘s are fully open and the VBV‘s are fully closed. The opposite occurs during power reductions. Modulation Schedule Revisions The EEC increases the compressor stall margin during rapid decelerations. (throttle chop) and reverse thrust operation. Rapid decelerations are sensed by the EEC. The large mass of the fan does not decelerate as quickly as the high pressure compressor. This causes an overload of airflow at the HPC inlet. To prevent a compressor stall, the EEC revises the normal VBV schedule so that the VBV‘s are open an additional 30 square inches. When the EEC senses that the decelerations of the fan and compressor have stabilized, it returns to the normal VBV schedule. During reverse thrust operation, the reversed fan air disturbs the airflow at the engine inlet. To ensure that the engine does not stall, the EEC revises the normal VBV schedule so that the VBV‘s are open an additional 30 square inches until reverse thrust is stopped. The VSV‘s are closed an additional four degrees from the normal schedule during reverse thrust.

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ENGINE AIR SYSTEM

Figure 12 SCL

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VSV and VBV Control Schedule Page: 25

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ENGINE AIR SYSTEM EICAS INDICATIONS Position indications appear on the EICAS EPCS page for the following engine air system components: -- Variable Stator Vane (VSV) Actuators -- Variable Bypass Valve (VBV) Actuators -- High Pressure Turbine Cooling (HPTC) Valve The indications are in percent of maximum angle, with 0 % equal to fully closed positions and 100 % equal to fully open. The ranges for the indications are from --5.0% to 105.0%. Parameter values are presented on the EICAS EPCS page for the following temperatures and pressures used to control engine air system components: -- Ambient (Static) Pressure (PO) -- HPC Discharge (Burner) Static Pressure (PS3) -- HPC Inlet Temperature (T25) -- HPC Discharge (Burner) Temperature (T3) The PO pressure indication range is from --1.5 to 20 PSIA. The PS3 indication range is from --5 to 600 PSIA. The T25 indication range is from --55 to 160 ° C. The T3 indication range is from --55 to 650 ° C.

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ENGINE AIR SYSTEM

Figure 13 SCL

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Engine Air System EICAS Indications Page: 27

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ENGINE AIR SYSTEM

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ENGINE AIR SYSTEM - OPERATION Core Compartment Cooling Valves The CCCV is closed at stabilized cruise power when the aircraft is above 17,000 Ft. altitude and the EGT is less than 699 ° C. Cooling airflow to engine accessories is reduced when the CCCV is closed. The CCCV fail--safe is open. HPTC Valve The HPTC valve opens at cruise power settings when the aircraft is above 17,000 Ft. altitude and N2 is between 82 % and 98 %. Turbine case cooling airflow is increased when the valve is open. The HPTC valve will fail--safe to closed. Variable Stator Vanes The VSV‘s modulate from fully closed during starting to fully open at takeoff power. The modulation schedule changes during reverse thrust operation. The VSV‘s fail--safe to closed.

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Variable Bypass Valves The VBV‘s modulate from fully open during starting to fully closed at takeoff power. The modulation schedule changes during rapid deceleration and reverse thrust operation. The VBV‘s fail--safe open.

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ENGINE AIR SYSTEM

Figure 14 SCL

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Engine Air System Operation Page: 29

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BEARINGS AND SUMPS

AIR EXTRACTIONS

General The two rotor systems contain 7 bearings wich are housed inside three oil sumps. the oil sumps are called A, B+C and D sump. A Sump houses bearing Nr 1 Ball N1 2 Roller N1 3 Roller N2 B+C Sump houses bearing Nr 4R Roller N2 4B Ball N2 5 Roller N2 D Sump houses bearing Nr 6 Roller N1

Fan Air Fan air is used for cooling: -- HPT and LPT Active Clearance Control -- Core Compartment Cooling -- Ignitor Leads -- IDG Oil Cooler

Sump sealing and pressurization All oil sumps are sealed with labyrinth type air and oil seals. Sump seal (Cavity) drains Drain lines are installed in the sump cavities (air pressurization chamber) to route leaking into : A sump radial drive shaft housing B+C sump LP recoup air exit D sump drain holes TRF struts (are not vented).

5th Stage LPC Sealing A, B+C and D sumps Cooling B+C sump, N1 rotor shaft and N2 compressor rotor. 7th Stage HPC Cooling -- HPT rotor (aft side) -- LPT rotor (fwd side) -- LPT 1st Stage Nozzle Guide Vanes (leading edge) 11th Stage Cooling -- HPT 2nd Stage Nozzle Guide Vanes and -- HPT 2nd Stage Stator Support.

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14th Stage HPC Cooling -- HPT 1st Stage Nozzle Guide Vanes -- HPT 1st and 2nd Stage rotor blades -- HPT rotor spool -- HPT stator case.

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Figure 15 SCL

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Sumps Airflow Page: 31

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RECOUP HP Recoup This is leaking CDP air wich has to be brought away from the compressor rear frame. Since the pressure and quantity are very high, the air is used for cooling purposes on the LPT 1st stage NGV trailing edge and then routed into the hot gas stream. LP Recoup Air out of the same part CRF, only lower on pressure, is routed via 3 tubes into the gas streamat the exhaust.

AIR EXTRACTION FOR AIRCRAFT USE 8th Stage HPC -- Pneumatic -- Thrust Reverser 11th Stage HPC Servo (muscle) pressure -- IDG air cooler valve -- Core Compartment Cooling Valve

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14th Stage HPC -- Pneumatic -- Thrust Reverser

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ENGINE AIR SYSTEM

Figure 16 SCL

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Airflow General Page: 33

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NOTES :

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ATA -- 76 ENGINE CONTROL SYSTEM TABLE OF CONTENT Engine Control Thrust lever Assy Autothrottle Clutch Pack Assy Thrust Lever Angle Resolver Fuel Control System Electronic Engine Control Control Alternator Fan T12 Sensor Compressor T25 Sensor EEC Discretes Circuit Card EEC Operation Power and Mode Select EEC Channel Reset EEC Control Mode EEC Engine Idle Select Hydromechanical Unit HMU Fuel Metering Operation Engine & Fuel Control EICAS Message

002 004 006 008 012 014 020 022 024 026 028 030 032 034 038 040 042 046

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ENGINE CONTROL SYSTEM

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ENGINE CONTROL SYSTEM General The General Electric CF6--80C2 is an engine which has a Full Authority Digital Electronic Control (FADEC) system, it is a computer--based engine control system. Each engine on the 767 has its own independent engine control system. The main component of the FADEC system is a control box called Electronic Engine Control (EEC). The FADEC system is divided into subsystems to perform two basic functions: -- information processing -- and engine control. The information processing functions receive, manipulate, and send large amounts of data. The EEC gets information about the environment and operating conditions within the engine. The EEC uses this information to control the engine. The EEC also sends data and messages to EICAS, the SEI and the PIMU. The engine control functions control the engine fuel and air systems to operate the engine efficiently at all rated performance levels. The engine systems that the EEC controls include fuel flow, primary engine airflow, turbine case cooling, parasitic (internal engine) cooling airflow, ignition, and engine starting systems. Fuel control is covered in this chapter. The other engine systems are covered in other chapters.

A dedicated control alternator generates power for the EEC when the engine is running. The EEC also receives aircraft power during engine start, EEC maintenance test, and as backup power. Fuel control switches in the flight compartment control a high pressure fuel shutoff valve in the HMU. This assures that the engine can be shutdown regardless of EEC inputs and failures. A microswitch pack is mechanically actuated by the thrust lever linkage. The microswitch pack acts as an interface between the thrust levers and other user systems. The EEC sends signals to EICAS and the Standby Engine Indicator (SEI) for indication. The EEC receives digital signals from the Thrust Management Computer (TMC) and Air Data Computer (ADC). The Flight Management Computer (FMC) is also linked to the EEC through the TMC. The EEC discretes card sends pneumatic demand signals to the TMC. The TMC sends these signals to both EEC‘s . The EEC discretes card sends an analog engine idle control signal directly to the EEC.

Fuel Control System Engine fuel flow is controlled by the EEC. There are no mechanical engine control connections between the flight compartment and the engines. The EEC must be operational for the engine to run. The EEC receives input signals from the thrust levers through thrust lever resolvers, (TRA) and from engine sensors. The EEC controls the hydromechanical unit (HMU) using analog electrical signals. The HMU controls thrust by controlling fuel flow to the fuel nozzles in the engine combustor.

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Figure 1 SCL

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Engine Controls Schematic Page: 3

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THRUST LEVER ASSEMBLY General The thrust lever assembly controls the amount of thrust and its direction (forward or reverse). The assembly is in the center control stand. The crank arm is connected to the reverse thrust lever assembly. The forward thrust lever assembly is linked to the crank arm with the thrust reverser latch. Forward / Reverse Thrust Lever Interlock The forward / reverse thrust lever interlock has a tab and a pawl that keep the forward and reverse thrust levers from operating at the same time. The pawl engages a tab to prevent lifting the reverse thrust lever unless the forward thrust lever is at idle. The pawl enters a slot in the structure when the reverse thrust lever is lifted to prevent advancing the forward thrust lever.

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Forward Thrust Operation When the reverse thrust lever is down (stowed), the thrust reverser latch is engaged. This links the forward thrust lever to the crank arm. The control rods move upward to increase forward thrust when the forward thrust levers are advanced. Reverse Thrust Operation When the reverse thrust lever is lifted, the thrust reverser latch disengages. This allows the crank arm to move the control rods downward to increase reverse thrust. The reverse idle detent assembly is a cam and a roller that give tactile feedback of reverse thrust lever position. In forward thrust, the cam and roller move together. In reverse thrust, the cam is stationary and the roller moves with the reverse thrust lever. The roller enters the cam detent to give a tactile indication that the reverser is commanded to deploy and that reverse thrust is at idle.

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ENGINE CONTROL SYSTEM

Figure 2 SCL

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Thrust Lever Assy Page: 5

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AUTOTHROTTLE CLUTCH PACK ASSEMBLY General The autothrottle clutch pack assembly is the interface between the autothrottle system and the engine fuel control system. It is located in the forward equipment center. The microswitch pack is linked to the clutch pack assembly through the forward cable drum. It is the interface to other aircraft systems. The switch pack is mounted below the drum. Autothrottle Clutch Packs The autothrottle clutch packs serve two purposes: -- to supply friction and feel’ for the thrust levers (manual), -- and to allow the autothrottle servo unit to move the thrust levers (automatic) . The clutch packs are mounted on a common shaft. The thrust levers are connected to one face of a clutch pack. The autothrottle servo unit is connected to the other face of both clutch packs. The clutch friction is set to supply the correct ”feel” when the thrust levers are moved manually against the autothrottle servo unit. When the autothrottle is engaged, the autothrottle servo unit moves the thrust levers through the clutch packs. The clutch packs make manual override of the servo unit possible at all times.

To adjust the switch group, place the thrust levers at the proper angle as described in the maintenance manual. A scale on the forward drum indicates the position. Push on the lock channel to disengage the adjusting bolt. Rotate the bolt to adjust the switch. Check that the position is correct by a continuity test on the appropriate pins in the electrical connector. When the position is correct, release the lock channel to re--engage the bolt. See Maintenance Manual Chapter 22--32 for details. Switch Titles -- S1 , S5 L / R LANDING WARNING. -- S2, S3 L AUTOBRAKE / AUTOBRAKE RTO. -- S6, S7 R AUTOBRAKE / AUTOBRAKE RTO. -- S8, S11 L / R THRUST REVERSER DCV. -- S1O, S14 L / R SPEEDBRAKE RETRACT. -- S12, S16 L / R TMS THRUST REVERSE. -- S17 LOAD SHED / PRESSURE CONTROL L. -- S18 LOAD SHED / PRESSURE CONTROL R.

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Microswitch Pack The microswitch pack has two cam following arms and two sets of switches for each engine. Cam surfaces machined on the lower half of the forward drums move the arms. This activates the switches to send thrust lever position signals to other aircraft systems. Switch Replacement and Adjustment The individual switches of the microswitch pack may be replaced, but the entire switch pack must first be removed. There is an adjustment screw for each microswitch. These screws are turned to get ah switches in the group to activate at the same time. This adjustment is best done on the bench before installation. In addition, there is an adjusting bolt for each group. The bolt is turned to get the switches to activate at the correct thrust lever angle.

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ENGINE CONTROL SYSTEM

Figure 3 SCL

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Autothrottle Clutch Pack Assy Page: 7

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

THRUST LEVER ANGLE (TLA) RESOLVERS The thrust levers control engine thrust. Each thrust lever is mechanically linked through the autothrottle clutchpack to a two channel thrust lever angle (TLA) resolver. The TLA resolver is a rotary transducer. The clutchpack turns the resolver rotor when the thrust lever is moved. The resolvers are mounted to the clutchpack assemblies in the forward equipment center. Access is through the forward equipment center access door. Each resolver has two sets of electrical outputs that are a function of the thrust lever angle. One signal from each resolver goes to EEC channel A, the other signal goes to EEC channel B. The TLA resolver rotor receives an ac--signal excitation from the active EEC channel. The excitation induces an ac--signal response in each of two coils that are mounted perpendicular to each other on the resolver stator. The phase angle difference between the two coil response signals varies as a function of thrust lever and resolver rotor position. The EEC senses this phase angle difference and uses it to determine commanded N1.

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ENGINE CONTROL SYSTEM

Figure 4 SCL

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Thrust Lever Angle Resolvers Page: 9

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

THRUST LEVER AND RESOLVER ANGLES The TLA resolver gives thrust lever angle position to the EEC. The TLA resolver stator is held by stationary components of the autothrottle clutch pack assembly. Moving the left (right) thrust lever causes the left (right) TLA resolver rotate. When the forward thrust levers are pulled back to the idle stop with the reverse thrust levers in the down position (thrust reverser stowed), the thrust resolver angle (TRA) value on the EICAS EPCS page must be between 33.7 and 34.1 degrees. If external test equipment is used to directly measure the TRA, the angle for the left TLA resolver must be between 33.7 and 34.1 degrees, and the angle for the right TLA resolver must be between 55.9 and 56.3 degrees. When the forward thrust levers are pushed forward to the maximum thrust position, the TRA value on the EICAS EPCS page must be between 85.0 and 88.5 degrees. if external test equipment is used to directly measure the TRA, the angle for the left TLA resolver must be between 85.0 and 88.5 degrees, and the angle for the right TLA resolver must be between 1.5 and 5.0 degrees. When the forward thrust levers are pulled back to the idle stop, and the reverse thrust levers are pulled up to tile maximum reverse thrust position, the TRA value on the EICAS EPCS page must be between 3.0 and 8.0 degrees. If external test equipment is used to directly measure the TRA, the angle for the left TLA resolver must be between 3.0 and 8.0 degrees, and the angle for the right TLA resolver must be between 82.0 and 87.0 degrees.

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ENGINE CONTROL SYSTEM

Figure 5 SCL

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Thrust Lever and Resolver Angles Page: 11

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

FUEL CONTROL SYSTEM The fuel control system directly controls engine thrust. The system is designed around a full authority, dual channel, digital electronic engine control (EEC). It is mounted on the fan case at the 8:30 position. Fuel flow is metered by the hydromechanical unit (HMU) mounted on the front right side of the accessory gearbox. In addition, the HMU supplies servo fuel for the operation of the engine air system. The HMU gets control signals from the EEC and the aircraft Other components of the fuel control system include the control alternator, electrical fan inlet temperature (T12) sensors and the T25 / P25 temperature / pressure sensor. The control alternator supplies power to the EEC and is driven by the accessory gearbox. There are two T12 sensors mounted on the forward edge of the fan case. The T25 / P25 sensor is mounted to the fan frame at the HPC inlet. The temperature signals are sent to the EEC for power management. An optional pressure signal goes to EEC condition monitoring circuits.

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ENGINE CONTROL SYSTEM

Figure 6 SCL

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Fuel Control Components Page: 13

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

ELECTRONIC ENGINE CONTROL The electronic engine unit (EEC) manages the following engine functions: -- Compressor airflow control (Chapter 75) -- Core compartment cooling (75) -- Turbine case cooling (75) -- Engine / aircraft interface (EICAS, TMC, etc.) (76) -- Power management in response to commanded thrust (76) -- Engine limit protection (76) -- Built--in testing (76) -- Fault detection (76) -- Engine status indications (77) -- Maintenance indications (77) -- Thrust reverser interlock and control (78) -- Start / ignition control (74/80) The EEC is a two channel (A and B), digital electronic microcomputer. It is mounted using vibration isolators on the left side of the fan case at the 8:30 position. There are fifteen electrical connectors on the front side of the unit, identified as J1 through J15. Engine wiring harnesses are color coded for easy identification. There are four connections for pressure probes on the bottom of the unit. The unit is cooled by natural convection. The EEC is designed to support a variety of engine / aircraft combinations and different thrust ratings. An engine rating plug on connector J14 programs the EEC for the desired application. The plug is attached to the engine fan case by a lanyard and remains with the engine if the EEC is changed. It must be connected to the EEC to dispatch the airplane. The EEC has two modes of operation: -- control -- and test. The EEC is normally in the control mode. It is in test mode if the airplane is on the ground, the fuel control switch is in CUTOFF, and the EEC ground test switch is in TEST.

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EEC

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EEC Location Page: 15

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ENGINE CONTROL SYSTEM EEC CONNECTORS Various airplane and engine systems communicate with the EEC and have redundant paths to the EEC channels (channel A and channel B) The 15 electrical connectors on the EEC are grouped by aircraft interfaces (J1--J6) , on--engine components (J7--J13) and EEC use (J14 and J15). The connectors are assigned as follows: Aircraft J1 -- Ignition exciter 1 dc power in / out; ch A ground handling bus power in. J2 -- Ignition exciter 2 dc power in / out; ch E ground handling bus power in. J3 -- Fuel on; starter air valve open; ch A reset, EEC fault, digital data bus (ADC, TMC) in / out, ch A TLA resolver in / out. J4 -- Single / dual igniters; idle select; hard reversionary mode; ch E reset, EEC fault, digital data bus (ADC, TMC) in / out, ch B TLA resolver in / out. J5 -- Aircraft type; engine position (L or R); ch A thrust reverser position. J6 -- TMC disconnect; operating mode select (control or test); ch B thrust reverser position

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

Engine J7 -- Black -- ch A. J8 -- Brown -- ch B. N2 sensor; ESCV solenoid, ESCV position switches; HMU J9 -- Red --ch A . J1O -- Orange -- ch B Control alternator; starter air valve; N1 sensor; T12. J11 -- Yellow -- ch A. J12 -- Green --ch B. T25; HPTC valve; VSV actuators; VBV actuators. J13 -- Blue -- ch A and ch B. T3; T49; T5; engine oil temperature sensor; Fuel flow transmitter. Electronic Engine Control (EEC). J14 -- Engine rating plug receptacle. J15 -- Engine identification plug receptacle.

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The engine rating plug (P14) and engine identification plug (P15) are captive to the engine by lanyards. The EEC contains rating tables for multiple ratings. The P14 rating plug determines the rating used by the EEC. This plug must be connected to the EEC to dispatch. The engine identification plug (P15) provides engine hardware information to the EEC, included in this information are the: -- N1 modifier level. -- EGT shunt value. -- Active clearance control schedules. After an EEC replacement on an engine, J15 is also used to enter the serial number of the engine into the memory of the EEC. The P15 plug is temporarily removed from J15. The cable from the programming tool is connected. (See AMM 73--21--15 / 201 for details) . When the serial number programming is completed, P15 is re--installed in J15. Pressure Inputs The EEC has pressure transducers and signal conditioning circuits. The pressures measured are: -- Ambient pressure (PO) -- Compressor discharge pressure (PS3) One transducer for each channel measures PO through a small hole in the EEC case. A tube for PS3 goes to the EEC. The two channels send data to each other on a crosstalk data bus.

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ENGINE CONTROL SYSTEM

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EEC Connectors Page: 17

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

EEC INPUTS / OUTPUTS The EEC gets analog input data from the engine and aircraft. It also receives digital input data and discrete inputs from the aircraft. The EEC uses power from a control alternator when the engine is running, and from the aircraft when the engine is not running. The EEC sends analog output signals to the hydromechanical unit (HMU), engine air systems, thrust reverser interlock, and start / ignition systems. The EEC sends digital signals to EICAS and the propulsion interface monitor unit (PIMU). The two EEC channels are redundant and independent. Each channel receives the same inputs. The system is designed so that no single failure causes the engine to stop running. The EEC includes extensive self--test and fault recovery features. When the EEC is on, it monitors all critical functions and inputs. If an input signal is faulty or missing, the EEC usually uses the value input to the other EEC channel. If that input is faulty or missing, the EEC often calculates an approximate value for the missing data. The EEC takes the following actions when input data is faulty or missing: -- Engine sensor data is used to backup the air data computer (ADC) TAT and PO values. -- The EEC calculates a mach number if MACH is not received from the ADC. -- Cross--channel data is used if Tl2 or PO sensor data is invalid. -- Comparisons are made between N1, N2, P3 or T25 sensor data inputs using cross--channel data. If sensor values disagree, the closest to an EEC calculated value is used; if both sensor values are lost or invalid, EEC calculated values are used. -- Comparisons are made between TLA data inputs using cross--channel data. If both inputs are lost or invalid, the last TLA value is used during takeoff; otherwise, the TLA is reduced to idle. -- The EEC calculates values for the HMU fuel metering valve, VSV actuator and VBV actuator if the position data is invalid or missing. -- The HPTC, LPTC, ESCV, CCC valves and the thrust reverser interlocks fail--safe to open or closed. -- The EEC uses 28 V dc aircraft power if power is not available from the proper control alternator.

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ENGINE CONTROL SYSTEM

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EEC Inputs / Outputs Page: 19

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CONTROL ALTERNATOR The control alternator is a two--winding, three--phase alternator that supplies electrical power to the EEC. It is mounted on the front left side of the accessory gearbox just outboard of the lube and scavenge pump. The alternator has two major components, the rotor and stator. The rotor is mounted on a stub shaft extending from the accessory gearbox. The shaft has flats milled on three sides. The rotor has permanent magnets arid is held on the shaft with a nut. The stator is bolted to the gearbox over the rotor. It has two separate three phase windings. Each set of windings supplies a three phase power signal to one of two connectors on the forward face of the stator. The inboard connector supplies power to EEC channel A. The outboard connector supplies power to EEC channel B. The control alternator meets all EEC power requirements when N2 increases above 11 %. It continues to meet the requirements until N2 decreases below 9 %. If one phase of either or both windings fails, the control alternator continues to meet all EEC power requirements if N2 is above 45 percent.

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ENGINE CONTROL SYSTEM

SCL

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May -- 2001

Page: 20

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 10 SCL

JGB

May -- 2001

Control Alternator Page: 21

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

FAN INLET TEMPERATURE (T12) SENSOR There are two T12 electrical fan inlet temperature sensors. Each supplies inlet temperature data to one of the EEC channels. The sensors are identical and are mounted on the forward edge of the fan case at the 2:00 and 10:00 positions. The sensing element in the sensors is a resistive thermal device (RTD) . It is constructed by wrapping a platinum wire around a ceramic core. The resistance of the platinum wire is directly proportional to the temperature of the inlet airflow. The RTD element is enclosed in an airfoil housing. The housing protects the element from physical damage. It also prevents water and ice from making contact with the element and interfering with the sensors ability to detect the true temperature of the inlet airflow. EEC channel A sends a 10 milliamp signal to the left (10:00) sensor. The voltage drop across the RTD element is measured by the EEC and corrected for ram air effects to determine the inlet air temperature. The right (2:00) sensor operates the same way with EEC channel B. The operational range of the T12 input to the EEC is from --130 to + 212 ° F (--90 to +100 ° C).

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Lufthansa LAN Technical Training

ENGINE CONTROL SYSTEM

SCL

JGB

May -- 2001

Page: 22

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 11 SCL

JGB

May -- 2001

Fan Inlet Temperature Sensor -- T12 Page: 23

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

COMPRESSOR INLET TEMPERATURE (T25) SENSOR The compressor inlet temperature (T25) sensor is part of the T25 / P25 temperature / pressure sensor. The T25 / P25 sensor is mounted on the fan frame at the 7:30 position between the no. 8 and no. 9 fan struts. The sensor has two separate temperature sensing elements, one for each EEC channel. The sensing elements are protected by an airfoil housing. The P25 pressure output is not used. The temperature sensing elements are resistive thermal devices (RTD). They are constructed by wrapping a platinum wire around a ceramic core. When mounted, the sensor airflow is inserted into the compressor inlet airflow. The resistance of the platinum wire is directly proportional to the temperature of the airflow. Each sensing element is connected to one of two electrical connectors on the body of the sensor. One connector is for EEC channel A, and the other connector is for EEC channel B. Each EEC channel sends a 10 milliamp current to a temperature sensing element. The EEC measures the voltage drop across the platinum wire and converts the voltage to a compressor inlet temperature value. The operating range of the T25 input to the EEC is from --130 to +392 ° F (--90 to +200 ° C).

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Lufthansa LAN Technical Training

ENGINE CONTROL SYSTEM

SCL

JGB

May -- 2001

Page: 24

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 12 SCL

JGB

May -- 2001

Compressor Inlet Temperature Sensor - T25 Page: 25

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

EEC DISCRETES PRINTED CIRCUIT CARD One EEC discretes printed circuit card serves both engines. It is an interface between various pneumatic user systems and the TMC and FMC. The TMC supplies both EEC‘s with bleed state information. The card also supplies a time--delay for the idle select control circuits. The card is in the P50 card file in the main equipment center. Relays on the card connect inputs and outputs. The card has two sections, one for each engine. The 28 V dc battery bus and the left 28 volt dc bus supply power to the card’s left engine section. The 28 V dc battery bus and the right 28 V dc bus supplies power to the card’s right engine section. THE CARD IS STATIC SENSITIVE. DO NOT HANDLE BEFORE READING THE PROCEDURE FOR HANDLING ELECTROSTATIC DISCHARGE SENSITIVE DEVICES (REF 20--41--01). THE CARD CONTAINS DEVICES THAT CAN BE DAMAGED BY STATIC DISCHARGE.

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CAUTION:

SCL

JGB

May -- 2001

Page: 26

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

For Training Purposes Only

ENGINE CONTROL SYSTEM

Figure 13 SCL

JGB

May -- 2001

EEC Discretes Printed Circuit Card Page: 27

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

EEC OPERATION The two EEC channels (A and B) are identical and equally capable of controlling the engine. Each channel contains a power supply, central processor unit, digital interface unit, signal conditioning unit, data interface unit, and solenoid driver unit. The channels are physically separated within the EEC. The internal power supply for each EEC channel gets three--phase ac power from separate windings of the control alternator when the engine is running (N2 greater than 11 %). Aircraft power is supplied when: -- the engine is being started -- the engine fuel control switch is in the RUN position -- or the EEC maintenance engine power switch ( P61) is in the TEST position. Normally, aircraft power is used for ignition, pneumatic starter control valve operation, and power for some of the internal EEC solenoid drivers. Control alternator power is used for all other EEC functions. If both channels are healthy (no faults), the channel in control of the engine (active channel) switches with every engine start. If one or both channels have faults, the healthiest channel is always selected as the active channel during engine starting. If a fault is detected in the active channel during engine run, the standby channel takes control if it is healthier than the other channel. If both channels have faults, the channel with the least severe fault(s) takes control. If both channels have failed the engine is shut down. Detected faults are stored in the volatile memory of each channel. Fault information is shared between the two channels through the crosstalk data bus. Pressure transducers and signal conditioners for pressure inputs are located inside the EEC. There are separate pressure sensor circuits for each channel. When the engine is running, both channels have power, receive input signals, process data, and send information to aircraft systems and to the other EEC channel. However, only the active channel operates the servo valves, solenoids and relays to control the engine. Similar outputs from the standby channel are terminated inside the EEC by switching relays.

SCL

JGB

May -- 2001

Page: 28

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 14 SCL

JGB

May -- 2001

EEC Operation Page: 29

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

POWER AND MODE SELECT Power The EEC gets power from the aircraft during engine start, EEC test, and when the fuel control switch is in RUN. Aircraft power is used if power from the control alternator is not available, or when N2 is less than 11 % Each EEC channel has an independent power relay. The relays are energized through the start relay, the EEC maintenance test switch, or the channel reset relays when the fuel control switch is set to RUN.

Test Setting the EEC maintenance test switch on the P6 panel to TEST starts an EEC test. Power is supplied to the EEC and the EEC common return is connected to the ground test enable input of both EEC channels. During the test, all EICAS engine parameters that normally appear when the engine is running are shown.

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Mode Select If the EEC fails to receive a valid total pressure value from either ADC, the EEC operates in a soft reversionary control mode. If N2 is greater than 50 % , as sensed by the N2 speed card, the ALTN light in the EEC control switch comes on after 10 seconds and the EICAS level C message L (R) ENG EEC MODE appears. This message is also latched as an EICAS status and maintenance message. Operating one engine using the soft reversionary control mode can cause thrust lever stagger, depending on ambient conditions. To eliminate this, the flight crew can command the EEC to operate in a hard reversionary control mode. This is done by pressing the EEC control switch on the P5 panel. The EEC common return is connected to the mode select input when the EEC control switch is cycled from the normal to the alternate position. This tells the EEC that the hard reversionary control mode has been selected. In this mode, the ALTN light in the EEC control switch is on. The EICAS message L (R) ENG EEC MODE appears as a level C message and as latched status and maintenance messages. If N1 command is greater than N1 maximum by more than 2% when the EEC is in either reversionary control mode, the level B EICAS message L (R) ENG LIM PROT appears.

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May -- 2001

Page: 30

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 15 SCL

JGB

May -- 2001

EEC Power and Mode Selector Page: 31

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

CHANNEL RESET AND FUEL ON Channel Reset The channel reset signal causes the EEC to alternate the active channel between channel A and channel B. Both EEC channels get a reset signal through the reset relays when the fuel control switch is moved to CUTOFF. Channel A also gets a reset signal if the fire switch is pulled. It a channel reset signal is receive while channel A is the active channel, channel B will become the new active channel if it is at least as healthy as channel A. If channel A is healthier than channel B, channel A will remain as active channel.

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Fuel On When the fuel control switch is set to RUN and the fire switch is set to NORM, a fuel--on signal is sent to both EEC channels.

SCL

JGB

May -- 2001

Page: 32

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 16 SCL

JGB

May -- 2001

EEC Channel Reset and Fuel On Page: 33

Lufthansa LAN Technical Training

ENGINE CONTROL SYSTEM EEC CONTROL MODES General The EEC uses total air temperature (T2), ambient pressure (PO), and total pressure (PT2) to compute the N1 command needed to meet commanded thrust. The thrust rating logic uses N1 command and several EEC control systems to determine required fuel flow. Normal Control Mode The air data computers (ADC‘s) supply T2, PO and PT2 to each EEC. The left ADC sends data to channel A. The right ADC sends data to channel B. Engine temperature sensors send air data to the EEC. The left T12 sensor data goes to channel A. The right T12 sensor data goes to channel B. Each EEC channel has a PO input. Using the crosstalk data bus, the data from both ADC‘s, both T12 sensors, and both PO inputs are available to each channel. Each EEC channel compares the total air temperature inputs (T2 L ADC, T2 R ADC, T12 CH A, and T12 CH B) to select a T2 value for calculating N1 l command. The ambient pressure inputs (PO L ADC, PO R ADC, PO CH A, and PO CH B) are used to select a PO value. A PT2 value is selected by comparing total pressure inputs (PT2 L ADC and PT2 R ADC). The selected PT2 value is used to calculate mach number (Mn) , impact Pressure (Q), the difference between ambient and standard day temperature (DTAMB), and the ambient temperature (TAMB). These values are used with T2 and PO to determine N1 command. The thrust lever angle (TLA) and bleed value received from the FMC are also used.

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Soft Reversionary Control Mode The normal control mode is used it PT2 L ADC and PT2 R ADC are both available and valid, and agree within 0.437 psia. Probe heat must also be ON. If these conditions are not met, the EEC automatically enters a soft reversionary control mode. If N2 is greater than 50 percent when the EEC switches to the soft reversionary control mode, the ALTN light on the EEC switch comes on, and the EICAS level C message L (R) ENG EEC MODE appears. The most recent DTAMB value while in the normal control mode is used for the soft reversionary control mode. This permits a smooth transition from the normal to soft reversionary modes. The fixed DTAMB value is used to calculate an assumed TAMB as altitude changes, and to calculate Mn and Q. N1 command is calculated using the assumed values for Mn, Q, TAMB, and DTAMB, and the PO, T2, TLA and bleed values. SCL

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May -- 2001

If the conditions required for normal control mode operation return while the EEC is in the soft reversionary control mode, the EEC goes back to the normal control mode if the current calculated Mn is within 0.1 of the current actual Mn. This ensures that control mode change does not cause significant changes in N1. Hard Reversionary Control Mode If an EEC remains in a soft reversionary control mode for an extended time, the two engines will develop different thrust levels. The hard reversionary control mode permits engine operation for extended periods. Manually selecting this mode ensures that both engines supply the same thrust at the same TLA position This mode is selected by pressing both EEC switches are pressed, the ALTN lights on the EEC switches come on, and the EICAS level C messages L ENG EEC MODE and R ENG EEC MODE appear. In the hard reversionary control mode, the DTAMB value used in calculating N1 command corresponds to the cornerpoint DTAMB value. The thrust can increase by using the cornerpoint DTAMB value instead of the DTAMB value used in the soft reversionary control mode. This can cause overboosting of the engine depending on actual ambient conditions and thrust lever angle.To prevent overboosting, the thrust levers must be pulled back to an intermediate position prior to selecting the hard reversionary control mode. The cornerpoint DTAMB value is used to calculate an assumed DTAMB as altitude changes, and to calculate Mn and Q. N1 command is calculated using the calculated values for Mn, Q, DTAMB, and DTAMB, and the PO, T2, TLA and bleed values.

Page: 34

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 17 SCL

JGB

May -- 2001

EEC Control Modes Page: 35

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

Limit Protection The EEC limits N1, N2, and the compressor discharge pressure (PS3). If any of the limits are approached or exceeded, the EEC reduces the fuel flow regardless of the TLA position. The N1 limit is 3,854 rpm (117.5%), the N2 limit is 11,055 rpm (112.5%), and PS3 is limited to 430 psid. The N2 limit schedule is used in addition to a mechanical overspeed governor in the hydromechanical unit (HMU). Acceleration / Deceleration Control The EEC limits the N1 and N2 acceleration and deceleration rates. If the commanded thrust increase is higher than allowable, the EEC limits fuel flow to the maximum rate allowed to prevent engine overboosting. If the commanded thrust decrease is lower than allowable, the EEC maintains a fuel flow sufficient to prevent engine flameout. This control ensures that all engines respond to thrust lever angle changes at the same rate.

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Idle Control The idle control calculates N2 demand. If minimum idle is not selected, the EEC calculates a flight idle N2 demand valve based on ambient temperature and pressure. When minimum idle is selected, the flight idle N2 demand is set to 6,050 rpm (61.6 %). The fuel flow is set to keep N2 speed at or above the flight idle N2 demand. If the N2 demand makes the compressor discharge pressure to low to meet bleed requirements, fuel flow is increased. Reverse Control Reverse control is active whenever the thrust reverser is not fully stowed. The EEC calculates the reverse thrust demand based on the thrust lever position. If the calculated reverse thrust N1 demand is greater than 3,280 rpm, or if the thrust demand is calculated to be greater than about 30,700 pounds, the fuel flow is reduced to ensure that these limits are not exceeded.

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JGB

May -- 2001

Page: 36

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 18 SCL

JGB

May -- 2001

EEC Control Modes Page: 37

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

EEC ENGINE IDLE SELECT CONTROL The engine operates at one of two idle speeds: -- minimum idle -- or approach (high) idle Minimum idle is generally used in the air. It is also used on the ground to reduce idle thrust while in the forward thrust mode. Approach idle is used during landing approach (flaps down) to meet the engine response time limits required for certification. To ensure an adequate flameout margin, approach idle is also used in flight when thermal anti--ice is on. The EEC sets the engine idle based on a signal loop between the EEC common return and the minimum idle terminals. if there is a signal loop, the EEC sets minimum idle. If the loop is broken, approach idle is set. Approach idle is the default setting. The EEC is commanded to approach (high) idle for any of the following: -- The thrust reverser PRSOV is energized. -- The thrust reverser is commanded to deploy and the fire handle is down in the normal position. -- The aircraft is in flight with flaps down (landing position). -- The aircraft is in flight with the thermal anti--ice system on. -- The aircraft is in flight with continuous ignition selected. Unless the EEC is commanded to approach idle for another reason, the EEC is commanded to change from approach idle to minimum idle: -- Five seconds after the flaps are raised past 23 ° after having been below 23 ° . -- Five seconds after the thermal anti--ice system is turned off after having been on. -- Five seconds after the aircraft has landed unless thrust reverser deployment is commanded. -- Immediately after power is removed from the T/R PRSOV and the reverse thrust lever has been stowed. If the idle commands to the both EEC channels do not agree, an EICAS message appears. Disagreements occur due to a faulty relay or idle command differences. The EICAS message IDLE DISAGREE appears as a level C message and as a latched maintenance message on the ECS / MSG page. SCL

JGB

May -- 2001

If the EEC senses that N1 is less than approach idle when the thermal anti--ice system is on, the EICAS message L (R) ENG LOW IDLE appears as a level C message and as a latched maintenance message. FADEC engines are susceptible to flame out at minimum idle when encountering inclement weather. The ignition select switch is used to command approach idle preventing possible flame out.

Page: 38

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 19 SCL

JGB

May -- 2001

EEC Idle Select Control Page: 39

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

HYDROMECHANICAL UNIT The fuel metering subsystem is completely contained in the hydromechanical unit (HMU). The HMU is mounted on the front, right side of the accessory gearbox. It is driven by a mechanical connection to the gearbox. The HMU responds to electrical signals from the EEC to meter fuel flow for combustion and to modulate servo fuel flow to operate the engine air systems. The HMU also receives signals from the aircraft fuel control system to control an internal high pressure fuel shutoff valve (HPSOV). Access to the HMU is through the right thrust reverser half. There are four external electrical connectors for electrical interfaces with the aircraft and EEC. Four fuel ports connect the HMU with the fuel pump and fuel nozzles. There are five hydraulic connections for control interfaces with the engine fuel and air systems. Each hydraulic interface is controlled by an electro hydraulic servo valve (EHSV) that varies servo fuel pressure in response to EEC signals. The fuel connections to the HMU are: -- Fuel inlet from the fuel pump -- Fuel discharge to the fuel nozzles -- Fuel bypass discharge to the fuel pump -- Servo fuel inlet from the servo fuel heater The hydraulic connections from the HMU are: -- Servo fuel pressure to the low pressure turbine cooling (LPTC) valve -- Servo fuel pressure to the high pressure turbine cooling )HPTC) valve -- Servo fuel reference pressure to the LPTC and HPTC valves -- Servo fuel pressure to the variable bypass valves (VBV‘s) -- Servo fuel pressure to the variable stator vanes (VSV‘s) The electrical connections to the HMU are: -- Fuel control signals from EEC channel A -- Fuel control signals from EEC channel B -- HPSOV solenoid inputs from the fuel control valves -- HPSOV position indication outputs to the EEC.

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May -- 2001

Page: 40

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

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ENGINE CONTROL SYSTEM

Figure 20 SCL

JGB

May -- 2001

Hydromechanical Unit Page: 41

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

HMU FUEL METERING OPERATION General The HMU has three hydraulic circuits: -- a fuel metering circuit -- a bypass circuit -- a servo control circuit. The fuel metering circuit controls fuel flow to the fuel nozzles in the engine combustor. It has a fuel metering valve and a high pressure fuel shutoff valve (HPSOV). Unentered fuel from the fuel pump goes to the FMV. Metered fuel from the FMV goes to the HPSOV. If the HPSOV is open, metered fuel is routed to the fuel nozzles. The bypass circuit is composed of a bypass valve, a differential pressure (delta P) regulator, and an overspeed governor. The fuel pump supplies more fuel than needed for the metered fuel flow. The bypass circuit returns excess fuel to the fuel pump. The servo control circuit divides the fuel supply from the servo fuel heater into regulated and unregulated servo flows. These flows operate actuators located both inside and outside of the HMU. The circuit has a servo regulating and distribution section and five electro--magnetic servo valves. One of these servo valves supplies servo pressure for FMV control and is discussed below. The other servo valves control pressure to engine air system actuators and are discussed under ENGINE AIR. Fuel Metering Valve A fuel metering valve (FMV) inside the HMU controls fuel flow to the nozzles. The hydraulically driven metering valve is controlled by the fuel metering valve EHSV. The EHSV has two coils, one for each EEC channel. The controlling EEC channel increases current through its EHSV coil to hydraulically open the FMV. If neither coil has power, the FMV closes. The FMV has two position indicating resolvers. One resolver is excited by, and provides a position feedback signal to, EEC channel A. The other resolver goes to EEC channel B.

SCL

JGB

May -- 2001

High Pressure Fuel Shutoff Valve A solenoid controls the position of the high pressure fuel shutoff valve (HPSOV). The fuel control switch and engine fire switch on the P10 panel control the HPSOV solenoid. The solenoid gets power directly from the 28 V dc from the battery bus. It has two latching coils: -- run -- and cutoff. Placing the fuel control switch to RUN energizes the run coil of the HPSOV solenoid. Placing the fuel control switch to CUTOFF, or pulling the engine fire switch, energizes the cutoff coil of the HPSOV solenoid. The solenoid is magnetically latched in the last commanded position. When the HPSOV solenoid is in the cutoff position, the HPSOV sends high pressure servo fuel to the pressurizing and shutoff valve to stop metered fuel flow to the fuel nozzles. When the solenoid is in the run position, the high pressure servo fuel is cutoff and the pressurizing and shutoff valve can open. When the pressurizing and shutoff valve is closed, a permanent magnet mounted to a translating structure on the valve is in close proximity with three reed--type switches. The magnet closes the three switches. One of the switch outputs goes to EEC channel A, one to EEC channel B, and one to the ENG VALVE disagreement light circuit. The EICAS level C message L (R) ENG FUEL VAL appears if the pressurizing and shutoff valve actual and commanded positions disagree. The ENG VALVE light on the P10 panel also comes on when the valve actual and commanded positions disagree. Bypass Valve The bypass valve has a piston inside a multi--ported sleeve. Unmetered fuel from the fuel pump enters the sleeve, is blocked by the piston, and is forced out of the sleeve ports. The fuel flow rate to the FMV, and the bypass return flow to the fuel pump, are controlled by moving the piston in and out of the sleeve varying the number of outlet ports. The piston position is controlled by the delta P regulator.

Page: 42

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BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

For Training Purposes Only

ENGINE CONTROL SYSTEM

Figure 21 SCL

JGB

May -- 2001

HMU Schematic Page: 43

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

Delta P Regulator The delta P regulator maintains a constant pressure drop across the FMV. This makes the fuel flow rate vary with the FMV position. The fuel flow rate is between zero and 30,000 pounds per hour. The regulator monitors the pressure difference between the unmetered fuel input and metered fuel output developed across the FMV. The regulator positions the bypass valve to equalize the two fuel pressures. If the FMV input pressure increases above the output pressure, the delta P regulator opens the bypass valve to increase bypass fuel flow to the fuel pump. If the FMV input pressure decreases below the output pressure, the bypass valve closes to decrease bypass fuel flow.

For Training Purposes Only

Overspeed Governor The overspeed governor senses N2 speed through the HMU mechanical drive from the accessory gearbox. If N2 exceeds 113.4 %, the governor overrides the delta P regulator input to the bypass valve to reduce metered fuel flow regardless of the FMV position. When the overspeed governor operates, it closes an overspeed indication switch inside the HMU. This switch is connected to the EEC. When the switch closes, the latched EICAS status and maintenance message L (R) ENG O/S GOV appears. When the engine is started, remaining fuel between the spar valve and the pressurizing and shutoff valve causes the overspeed governor to operate, closing the overspeed switch. The overspeed governor returns to normal operation at 50 % of N2. This performs a functional test of the overspeed governor. If the switch does not close during engine start, the L (R) ENG O/S GOV message appears.

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May -- 2001

Page: 44

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ENGINE CONTROL SYSTEM

Figure 22 SCL

JGB

May -- 2001

HMU Operation Page: 45

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ENGINE CONTROL SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

ENGINE AND FUEL CONTROL EICAS MESSAGES General The EEC monitors itself and the operation of the engine. When an internal, input, or output fault is found, the fault is stored in the EEC volatile memory. The EEC sends signals to EICAS for indication. Faults are transferred to the propulsion interface monitor unit (PIMU) non--volatile memory immediately after the aircraft has landed. EICAS Alert Messages The following alert messages for each engine appear on the EICAS primary engine parameters page: -- L (R) ENG LIM PROT is a level B message. It means that the EEC is in a revisionary mode and that the N1 thrust setting exceeds the maximum rating by 2 %. -- L (R) ENG SHUTDOWN is a level B message. It means that the engine fire switch has been pulled or the fuel control switch is in CUTOFF. There is no master caution light or aural warning. Other engine related messages are inhibited for 20 seconds. -- L (R> ENG CONTROL is a level C message. It means that the EEC is in a NO dispatch configuration. This message only appears when the aircraft corrected airspeed is below 80 knots. It occurs if both of the EEC channels are incapable of controlling the engine. The HMU fuel metering valve goes to the minimum idle stop. -- L (R> ENG EEC MODE is a level C message. It means that the engine EEC is operating in a reversionary mode. The message appears 5 seconds after the EEC starts operating in a reversionary mode. -- L (R) ENG FUEL VAL is a level C message. It means that the HMU high pressure fuel shutoff valve (HPSOV) actual and commanded positions disagree. The message appears if the disagreement exists for more than 6 seconds. -- L R) ENG LOW IDLE is a level C message. It means that the engine is at minimum idle with the flaps down or with the thermal anti--ice system on. The message appears if the condition exists for more than 6 seconds.

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May -- 2001

-- L (R) ENG RPM LIM is a level C message. It means that the EEC is limiting thrust due to N1 overspeed, and that additional thrust is not available. The message appears 3 seconds after the EEC starts limiting thrust. -- IDLE DISAGREE is a level C message. It means that one engine is at ”approach” idle while the other engine is at ’minimum’ idle. The message appears if the idle disagreement exists for more than 6 seconds. EICAS Status and Maintenance Messages Many EICAS status and maintenance messages relate to engine, HMU and EEC operation. In general, all of the messages indicate that the EEC is operating in a reduced capacity. They do not necessarily mean that the EEC is inoperative, but they do mean that the EEC may not be able to perform all of its normal functions. The following status and maintenance messages associated with engine control and aircraft dispatchability appear on the EICAS status or ECS / MSG pages: -- L (R) ENG EEC C1 is a status and maintenance message. It means that the EEC is in a time--limited dispatch configuration. In this condition, the aircraft can be dispatched. The problem must be corrected as required by GE engine type certificate data sheet number E13NE, note 18. This message is latched. -- L (R) ENG EEC C2 is a latched maintenance message. It means that the EEC is in a long time limited dispatch configuration condition. In this condition, the aircraft can be dispatched. The problem must be corrected as required by GE engine type certificate data sheet number E13NE, note 18. -- L (R) ENG O/S GOV is a status and maintenance message. It means that the HMU N2 overspeed governor has failed an initialization test. This message appears 5 seconds after the test failure and is latched. The following alert messages are also status and / or maintenance messages: -- L (R) ENG CONTROL is a latched status and maintenance message. -- L (R> ENG EEC MODE is a latched maintenance message. -- L (R) ENG LOW IDLE is a latched maintenance message. -- IDLE DISAGREE is a latched maintenance message. Page: 46

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ENGINE CONTROL SYSTEM

Figure 23 SCL

JGB

May -- 2001

Engine and Fuel Control Messages Page: 47

BOEING -- 767 / 300 CF6 -- 80C2 76 -- 00

NOTES :

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ENGINE CONTROL SYSTEM

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May -- 2001

Page: 48

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ENGINE INDICATING SYSTEM

BOEING -- 767 / 300 CF6 -- 80C2 77 -- 00

ATA -- 77 INDICATING SYSTEM TABLE OF CONTENT Engine Indicating Introduction Engine Indications Tachometer System N1 Shaft Speed N2 Shaft Speed N2 Speed Cards Tachometer System Operation EGT Indicating System EGT EICAS Indications Condition Monitoring Fan Discharge Pressure Ps14 LPT Inlet Pressure P49 LPT Discharge Temperature T5 Compressor Inlet Temperature P25 Standby Engine Indicator ( SEI ) Airborne Vibration Monitoring ( AVM )

002 006 008 010 012 014 016 020 028 030 032 034 036 038 040 044

Propulsion Interface Monitoring Unit ( PIMU ) EEC Faul Monitoring in Flight PIMU Fault Recording PIMU BITE Recent Flight EEC Fault Monitoring in Ground PIMU BITE Ground Test PIMU Maintenance Recall EICAS -- EPCS Pages

052 054 056 058 060 062 064 066

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ENGINE INDICATING Introduction The engine indicating system measures engine control and status parameters, and provides the parameters to aircraft systems for indication. Engine control parameters are measured by engine mounted tachometers, temperature sensors and pressure probes. Engine status parameters are provided by LRU‘s in the engine fuel, oil, air and thrust reverser systems. Engine indicating system data and messages are shown in the flight deck and main equipment center. Engine operating and status parameters are presented on the EICAS primary and secondary engine parameter displays, and on the PERF / APU, ENG EXCD and EPCS maintenance pages. Critical engine parameters are also shown on the standby engine indicator ( SEI ) if the EICAS system fails or is not powered. Messages related to engine performance appear on the EICAS status page and ECS / MSG maintenance page, and on the propulsion interface monitor unit ( PIMU ) in the main equipment center.

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ENGINE INDICATING SYSTEM General Description The engine indicating system provides flight deck display of engine parameters. The engine indicating system includes the following: -- Engine tachometer system: provides thrust indication to the flight deck and input to other Systems of rotor shaft speed (N1 and N2). -- Exhaust gas temperature indication: provides indication for crew monitoring and EGT input to the EEC. -- Airborne vibration monitoring system: measures the vibration of the engine. -- Engine N2 speed card: provides engine speed status to other aircraft systems. -- Standby engine indicator: provides backup indication for N1, EGT and N2 in the event of an EICAS failure. -- Propulsion interface monitor unit: stores fault data from the EEC and provides engine to airframe component communication.

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ENGINE INDICATING SYSTEM Engine Tachometer System There are two engine tachometer indications. The low pressure (fan) shaft speed which is called N1 and the high pressure shaft speed which is called N2.

Propulsion Interface Monitor Unit ( PIMU ) EEC internal, input and output faults are stored in volatile memory during flight. When the aircraft lands, the fault data is transferred to nonvolatile memory ( NVM ) in the PIMU for use during maintenance.

N1 Indication The primary thrust indication is N1. The N1 fan shaft speed sensor on the fan frame provides analog signals to the EEC to be converted to digital data and sent to EICAS and SEI. Separate analog N1 signals are also sent directly to the AVM signal conditioner and to EICAS for backup. The data is processed by EICAS and sent to the upper EICAS display for indication. N2 Indication The N2 shaft speed sensor provides an N2 output signal to the EEC, the N2 speed card, the AVM, and EICAS. The N2 output signal to the EEC is converted to digital data and sent to EICAS and SEI. N2 is processed and sent to the EICAS display for indication. The N2 speed card is designed to provide interface between the engine N2 speed sensor and various other systems on the airplane requiring a discrete signal of engine speed.

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Exhaust Gas Temperature ( EGT ) Indication The EGT system senses the internal gas temperature of the engine between the high and low pressure turbines. Eight EGT probes provide an output signal to the EEC where it is converted to digital data and sent to EICAS and SEI. The data is processed and sent to the upper EICAS display for indication. Airborne Vibration Monitoring ( AVM ) System The AVM system senses engine vibration levels and processes signals for EICAS display. There are two sensors. The No. 1 bearing accelerometer senses fan vibration. The compressor rear frame (CRF) accelerometer senses N2 rotor (core) vibration. The accelerometers provide vibration signals to the AVM signal conditioner. These signals are processed along with N1 and N2 signals by the AVM signal conditioner and are then sent to EICAS. The lower EICAS display provides vibration indication.

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Engine Indicating System Schematic Page: 7

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ENGINE TACHOMETER SYSTEM The engine tachometer system senses the speed of both engine rotor shafts ( N1 and N2 ) and sends N1 and N2 analog speed signals to EICAS, the EEC, and AVM. The system also sends an N2 analog speed signal to the N2 speed card. The signals are used for indication and control. The EEC sends digital N1 and N2 signals to EICAS and the SEI. EICAS uses the digital N1 and N2 signals for indication if either digital signal is available, and uses the analog signal for backup and comparison. The sensors are induction--type tachometers. The tip on each sensor has a permanent magnet with three coil assemblies. Each assembly has three separate circuits which send separate N1 and N2 speed signals to each channel of the EEC and to the airplane indicators.

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N1 FAN SHAFT SPEED SENSOR The N1 speed sensor uses signal pulses to measure N1 rotor speed. The sensor is mounted on the fan frame strut at the 2:00 position, just aft of the No. 3 strut. The sensor is held in place by two bolts and is accessible with the right T/ R half open. The N1 sensor consists of a stainless steel housing with three sensor coils in the tip and two electrical connectors at the other end. The sensor is 20 inches long and 3/4 inches in diameter. As the fan shaft rotates, 38 ferromagnetic teeth pass by the N1 sensor tip, inducing electromagnetic pulses in the sensor coils. The pulse frequency is directly proportional to fan speed. One of the teeth is taller than the others to aid tracking vibration for balancing. THE CF68002 FADEC N1 SENSOR IS NOT INTERCHANGEABLE WITH THE NON--FADEC CF6--80C2 ENGINE N1 SENSOR. DAMAGE TO THE TEETH AND / OR SENSOR WILL OCCUR IF THE INCORRECT SENSOR IS INSTALLED.

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N1 Fan Shaft Speed Sensor Page: 11

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N2 CORE SHAFT SPEED SENSOR The N2 speed sensor uses signal pulses to measure N2 rotor speed. The sensor is mounted on the forward side of the accessory gearbox, inboard of the HMU and adjacent to the core motoring pad, the sensor is held in place by two bolts and is accessible with both T / R halves open. The N2 sensor contains three sensor coils in the tip and two electrical connectors at the other end. As the N2 core shaft rotates it drives the accessory gearbox through the horizontal drive shaft. The horizontal drive shaft rotates the starter drive shaft. An idler gear with 12 ferro--magnetic lugs is rotated by the starter drive shaft. The lugs pass by the N2 sensor tip, inducing electromagnetic pulses in the sensor coils. The pulse frequency is proportional to N2 core shaft speed.

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N2 SPEED CARDS General The N2 speed cards are the interface between the N2 speed sensors and other aircraft and engine systems that require N2 speed signals. Two cards, one for each engine, are located in the P50 electrical systems and card file in the main equipment center. The cards are printed circuit cards and have two separate channels. Comparators control relays within each channel that send speed signals to user systems. CAUTION:

WARNING:

MOVING ENGINE N2 DISCRETE PRINTED CIRCUIT CARD CH. 1 SWITCH TO TEST CAUSES PROBE HEAT POWER TO BE APPLIED. PHYSICAL CONTACT WITH PROBE BODY CAN CAUSE SEVERE BURNS.

STATIC SENSITIVE. DO NOT HANDLE BEFORE READING PROCEDURE FOR HANDLING ELECTROSTATIC SENSITIVE DEVICES (20--41--01). CONTAINS DEVICES THAT CAN BE DAMAGED BY STATIC DISCHARGES.

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Operation Each N2 speed card channel gets power from the 28 V dc battery bus. Each channel gets the N2 core shaft speed sensor output signal. The signal is converted to a speed value by the N2 speed card sensing logic. The N2 speed value is compared to set values by four comparators. When the N2 speed value is determined to be above a fixed comparator value, N2 speed card relays are energized. The relay states permit user systems to determine if the N2 speed is above or below set values. If the channel 1 50% comparator disagrees with the channel 2 52% comparator for more than 10 seconds, the EICAS status and maintenance message L (R) ENG SPEED CARD appears. This is a latched message. The message is inhibited when the standby bus does not have power. Test Functions Channel 1 has a non--monentary toggle--type test switch. Channel 2 has a monentary toggle--type test switch. The two test switches permit functional test of both channels of the card when the engines are not running. Activation of both test switches at the same time for longer than 10 seconds indicates proper function if no EICAS message appears. Activation of the channel 1 test switch alone causes the EICAS L (R> ENG SPEED CARD status message after 10 seconds to check that the two channels are properly functioning.

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N2 Speed Cards Page: 15

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ENGINE TACHOMETER SYSTEM OPERATION The coil assemblies in each sensor tip send analog signals to the EEC. One coil in each assembly sends a signal to channel A of the EEC, and another sends a signal to channel B. Two electrical connectors send the signals to the EEC channels on different wire bundles. The third coil in the N1 sensor sends signals to EICAS and AVM. The third coil in the N2 sensor sends signals to EICAS, AVM, and to the engine N2 speed cards. The EEC converts the analog signals to digital data. The EEC sends digital N1 and N2 data to EICAS and the SEI. An EICAS latched level S, M message L (R) ENG ANALOG N2 appears when the analog N2 input to EICAS is less than 40 % and the digital N2 input from the EEC is greater than idle for 10 seconds.

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Engine Tachometer System Operation Page: 17

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ENGINE TACHOMETER SYSTEM EICAS INDICATIONS EICAS -- Primary Engine Display Actual N1 for each engine appears on the EICAS primary engine display as a digital readout and as a pointer on a round analog scale. The round analog scale has a white arc with a red line limit. A double yellow line for the N1 maximum limit is calculated by the EEC based on current ambient air temperature and pressure, and pneumatic demand. If the output from both EEC channels is invalid, signals from the TMC are used to generate the yellow line. The N1 command sector shows the difference between actual N1 and commanded N1. The EEC gets commanded N1 from the thrust lever angle (TRA) resolver. The actual N1 speed pointer sweeps off the command sector as speed changes. When the engine speed is stable, there is no command sector. Actual N1 digital readout and the enclosing box appear in white. The digits, box, and analog pointer change color from white to red when the red line limit is exceeded. During an exceedance, the scale extends to the pointer. The highest value of N1 exceedance appears in white digits under the N1 digital readout. The thrust reference cursor is calculated using signals from the FMC or, if the FMC is inoperative, from the TMC. The cursor is magenta in color when the FMC autopilot is engaged in VNAV mode. The cursor is green in color when the TMC is in control. The value of the thrust reference cursor appears in green above the N1 digital readout box. The thrust mode selected on the thrust mode select panel appears in green at the top of the display.

EICAS -- PERF / APU Page N1 command, N1 maximum, N1 actual and N2 actual appear in digital form on the PERF / APU maintenance page. EICAS -- Engine Exceedance Page The highest N1 and N2 exceedance values reached during engine operation appear in digital form on the engine exceedance maintenance page. The total time that N1 and N2 exceeded their red line limits also appears in digital form on the engine exceedance page.

EICAS -- Engine Secondary Display Actual N2 for each engine appears on the EICAS secondary engine display as digital readout and a pointer on a round analog scale. The round analog scale has a white arc with a red line limit. The actual N2 digital readout, box, and analog pointer change color from white to red when the red line limit is exceeded. During an exceedance, the scale extends to the pointer. The highest value of N2 exceedance reached appears directly under the N2 digital readout box in white numbers. A magenta fuel on command line appears when the engines are shutdown. The value is set at 15 % of N2 on the ground and 10 % of N2 in flight.

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Engine Tachometer System EICAS Indications Page: 19

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EXHAUST GAS TEMPERATURE (EGT) INDICATING SYSTEM General Description The EGT indicating system gives an indication of the average gas temperature at the LPT inlet of each engine. Eight EGT thermocouple probes are mounted in the high pressure turbine exhaust at engine station 4.9. An upper and a lower wiring harness join the probes to a junction box mounted on the left side of the engine. From the junction box, an overall chromel signal and an overall alumel signal are sent to EEC channels A and B. The EEC converts the signals to digital data and sends them to EICAS for indication.

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EGT THERMOCOUPLE (T49) PROBES General EGT alumel / chromel probes sense engine exhaust temperatures for flight deck indication and engine operation. The probes are connected to the EEC through a junction box. Each of the eight EGT probes senses the temperature of the gas flow between the HPT and LPT. The EGT probes are mounted around the LPT forward case at station 4.9, just forward of the low pressure turbine second--stage rotor blades. Characteristics Each probe has two parallel--wired thermocouple junctions. The junctions are at two different immersion depths within a protective sleeve.

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Removal and Installation Each probe is mounted with two bolts. An arrow inscribed in the top of the probe shows the correct orientation of the probe. The probes can be replaced individually. Each probe has exposed studs to permit continuity and resistance checks without removal. Thermocouple cables attach to studs on each thermocouple probe. The chromel lead goes to the small stud, and the alumel lead goes to the large stud.

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EGT THERMOCOUPLE CABLE AND JUNCTION BOX General The thermocouple cable connects the probes to a junction box on the engine. Junction Box The junction box conditions the EGT probe signals. The conditioning circuit averages the four chromel signals and the alumel signal from each harness into one overall chromel signal and one overall alumel signal. The junction box has an output connector that sends the conditioned signal to the EEC. The junction box is mounted on a bracket on the LPT cooling air tube near the HPC left horizontal splitline.

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Thermocouple Cable The thermocouple cable consists of an upper and lower cable harness. The upper harness connects probes 1, 2, 7 and 8. The lower harness connects probes 3, 4, 5 and 6. There is one common wire for all of the alumel studs. There is one wire for each individual chromel stud. The cables are mounted around the LPT forward case and are supported by brackets on the LPT and HPT case splitline. The forward portions of the thermocouple cables go along the left side of the HPC stator case and connect to the junction box.

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EGT Probe, Cable and Junction Box Page: 25

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EGT INDICATING SYSTEM OPERATION General description When the engine is running, hot gases from the high pressure turbine circulate around the probes. The hot gases heat the junction of the dissimilar metals (chromel and alumel) The difference in expansion rates between the two metals creates a voltage potential. A circuit is formed in the indicating system when the other ends of the leads are joined (the cold junction) in the EEC. The EEC processes the analog signal, sends it to both channels A and B, converts it to digital data and sends it to EICAS and the SEI.

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ENGINE INDICATING SYSTEM EGT INDICATING SYSTEM EICAS DISPLAYS The EGT appears on the EICAS primary display. The display has both digital and analog round dial EGT indications A digital EGT indication also appears on the PERF / APU page. EGT exceedance histories and profiles appear on the ENG EXCD page. When the EEC does not have power, there is no digital EGT input to EICAS. Since there is no analog backup, the EGT signal is not present when the EEC does not have power (engine shutdown) The primary display shows the analog round dial and box with no pointer or digits, and EGT digital indications are not shown on the PERF / APU page. EGT Indication -- EICAS Primary Display The EGT analog display has a white arc with yellow band and red line limit markers. A hot start limit marker appears when N2 is less than 50 %. A pointer shows actual EGT. Actual digital EGT appears in white numbers inside a white box. The numbers, box, and pointer turn yellow or red during yellow band or red line exceedances, respectively. The numbers, box, and pointer turn red during engine start if the EGT exceeds the hot start limit. This limit marker disappears after the engine idles at greater than 60 % of N2 for 10 seconds. The highest red EGT exceedance appears in white under the box.

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EGT Indication -- Engine Exceedance Page The engine exceedance page shows EGT exceedance histories and profiles for starting and continuous operation. An EGT red line exceedance begins when the EGT increases above the EGT red line limit (960 ° C), and ends when it decreases below the red line limit. During an EGT red line exceedance, EICAS records the time of exceedance and the highest EGT value reached. Following each EGT red line exceedance, EICAS adds the exceedance time to any previous time recorded in the exceedance nonvolatile memory, and records the maximum exceedance value if it is larger than the previously recorded value. The total EGT red line exceedance time and maximum red line exceedance value appear next to the EGT RED call out on the engine exceedance page. Left engine data is on the left, and right engine data is on the right. During engine starts, an EGT start exceedance begins when the EGT increases above the hot start limit (750 ° C) and ends when it drops below the hot start limit. Maximum EGT start exceedance values and total SCL

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exceedance times are recorded in a manner similar to the recording of EGT red line exceedance data. The total EGT start exceedance time and maximum start exceedance value appear next to the EGT START title on the engine exceedance page. Left engine data is on the left, and right engine data is on the right. EGT exceedance profiles are recorded by EICAS and appear at the bottom of the engine exceedance page. Left engine profiles are on the left, and right engine profiles are on the right. An exceedance profiles shows, for a specific exceedance event, the time that the EGT exceeded various temperatures. Up to 11 temperatures appear in an exceedance profile. An EGT AMBER exceedance profile is recorded when the EGT increases above the EGT amber band lower limit (925 ° C) but does not increase above the EGT red line limit (960 ° C). An EGT RED exceedance profile is recorded when the EGT increases above the EGT red line limit (960 ° C) . The lowest temperature for the EGT AMBER and EGT RED exceedance profiles is the EGT amber band lower limit (925 ° C) and the interval between temperatures is 10 ° C. An EGT START exceedance profile is recorded during EGT start exceedances. The lowest temperature for EGT START exceedance profiles is the hot start limit (750 ° C) and the interval between temperatures is 15 ° C. The maximum EGT value reached during the exceedance event appears next to the profile. EICAS records EGT exceedance profiles if the exceedance nonvolatile memory is clear, or if the priority of the new exceedance profile is equal to or greater than the priority of a previously recorded profile. EGT RED and EGT START exceedances each have the highest priority. EGT AMBER exceedances have the lowest priority. EICAS overwrites a previous EGT START or EGT AMBER exceedance profile with a new EGT RED exceedance profile. A new EGT START exceedance profile overwrites a previous EGT RED or EGT AMBER exceedance profile. A new EGT AMBER exceedance profile only overwrites a clear memory or a previous EGT AMBER exceedance profile.

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EGT Indicating EICAS Displays Page: 29

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CONDITION MONITORING SYSTEM The condition monitoring system includes three pressure probes and one temperature sensor which send analog signals to the EEC. The EEC converts the analog signals to digital data and sends a multiplexed signal to the PIMU. The ÁRINC communications and reporting system (ACARS) uses this information for diagnosis and fault isolation. The condition monitoring system includes signals from the following engine mounted sensors: -- Fan discharge pressure PS14 probe -- LPT inlet pressure P49 probe -- LPT discharge temperature T5 sensor -- Compressor inlet pressure P25 probe

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Condition Monitoring System Page: 31

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FAN DISCHARGE PRESSURE (PS14) PROBE The fan discharge pressure (PS14) probe senses static fan discharge pressure and sends the pressure signal to a PS14 transducer inside the EEC. The PS14 probe is mounted on the aft fan case just above the EEC at the 10:30 position. The sensor has a static pressure tap, mounting flange and pressure output port with a pressure tube that goes to the EEC. The operational range of the PS14 input to the EEC is between 2 and 30 psia.

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Fan Discharge Pressure Probe - Ps14 Page: 33

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LPT INLET PRESSURE (P49) PROBE The LPT inlet pressure (P49) probe senses the total pressure of the LPT inlet airflow. The P49 probe is mounted on the low pressure turbine case at the 3:30 position. The probe has a pressure tube that goes to a pressure transducer inside the EEC. Operational range of the P49 input to the EEC is between 25 and 120 psia.

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LPT Inlet Pressure Probe - P49 Page: 35

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LPT DISCHARGE TEMPERATURE (T5) SENSOR The T5 sensor measures the LPT discharge temperature It has two chromel--alumel type thermocouples and an electrical connector. The T5 sensor is mounted on the aft end of the turbine rear frame at the 9:30 position. The operational range of the T5 sensor input to the EEC is from --76 to +1571 ° F (--60 to +855 ° C).

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LPT Discharge Temperature Probe -- T5 Page: 37

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COMPRESSOR INLET PRESSURE (P25) PROBE The P25 probe is an integral part of the compressor inlet temperature / pressure T25/P25 sensor. The P25 probe senses the total pressure of the high pressure compressor inlet airflow. The T25 / P25 sensor is mounted on the fan frame hub outer surface at the 7:30; position. The P25 probe has a pitot tube for sensing pressure. The pressure signal goes to a P25 pressure transducer inside the EEC. The operational range of the P25 input to the EEC is from 2 to 75 psia.

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Compressor Inlet Pressure Probe -- P25 Page: 39

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STANDBY ENGINE INDICATOR (SEI) The SEI supplies backup N1, N2 and EGT indications when EICAS does not have power, or is not showing the primary engine parameters. The SEI is on the right side of the P1--3 panel. The SEI has eight LED digital displays. Six displays show N1, N2 and EGT for both engines. The SEI has its own power supply. There is a test switch to test the SEI for correct operation. There is a switch on the face of the SEI to select AUTO or ON. In AUTO the SEI display is inhibited if EICAS primary engine parameters are available. The SEI display is continuous in the ON position. The SEI receives N1, N2 and EGT data from the EEC. If the SEI is on, but the EEC does not have power (engine shutdown), N1, N2 and EGT indications do not appear on the SEI. THE WORDS FAIL NO LIMIT APPEAR ON THE FACE OF THE NEW SEI, IF THE SEI IS REPLACED AND THE OPERATIONAL PLACARDS FOR THE GE CF6--80C2F ENGINE, DO NOT REMOVED FROM THE OLD SEI AND INSTALLED ON THE NEW ONE BEFORE IT IS INSTALLED IN THE PANEL.

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ENGINE INDICATING SYSTEM

Figure 20 SCL

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Standby Engine Indicator Page: 41

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OPERATION AND TEST Operation and Displays Twenty--eight volt dc power is supplied from the battery bus. Digital signals from the EEC are received for display. When the select switch is in AUTO the display of engine data is controlled by the EICAS system SEI inhibit circuit. The engine parameters appear when the switch is in AUTO and EICAS is not showing primary engine parameters. This occurs when EICAS does not have power, has failed, or is in TEST mode. The engine parameters appear on the SEI when the select switch is in the ON position even if EICAS is displaying primary engine parameters.

Fault Codes -- 111 EPROM checksum failure -- 222 RAM failure -- 333 Frequency processing hardware -- 444 Input select hardware -- 555 Power supply -- 666 ARINC receiver failure (L EEC CHN A) -- 777 ARINC receiver failure (R EEC CHN A) -- 888 ARINC receiver failure (L EEC CHN B) -- 999 ARINC receiver failure (R EEC CHN B)

Fault Monitoring and Test Displays The SEI continuously monitors itself for correct operation. Zeros appear for invalid or missing input signal. A BITE test begins during: -- (1) airplane power up -- (2) when the SEI display comes on automatically -- (3) by selecting ON. No indications other than normal engine parameters appear during this test unless a fault is detected. If a fault occurs, dashes (-- -- -- ) appear for both N1 indications. Before replacing the SEI, use the T--switch to run a built--in--test to get specific fault codes. To run an SEI test, turn the T--switch clockwise with a small screwdriver. The SEI can only be tested if the SEI has power (display control switch is in the ON position or in the AUTO position if EICAS is not operative). The following fault codes can appear both N1 indicators during a test. These codes are used by the repair in shop.

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ENGINE INDICATING SYSTEM

Figure 21 SCL

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May -- 2001

SEI Operation and Test Page: 43

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ENGINE INDICATING SYSTEM

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AIRBORNE VIBRATION MONITORING (AVM) SYSTEM (ENDEVCO) General The airborne vibration monitoring (AVM) system continuously monitors engine vibration levels to detect engine malfunctions. Two accelerometers are mounted on each engine. There is one AVM signal conditioner. Four signals are sent from each engine to the AVM signal conditioner. They are: -- N1 speed from the N1 fan shaft speed sensor -- Fan vibration from the No. 1 bearing accelerometer -- N2 speed from the N2 speed sensor -- Core vibration from the CRF accelerometer The accelerometers sense vibration from N1 and N2 shafts. The signal conditioner uses these accelerometer vibration signals and N1 and N2 speed signals to determine the amplitude of the individual rotor vibrations for each engine. The signal conditioner sends the information to EICAS for indication, and for recording vibration data for fan trim balancing. Accelerometers The No. 1 bearing accelerometer is in the A sump on the No. 1 bearing housing. It is accessible only by major engine disassembly. The CRF accelerometer is on the forward side of the compressor rear frame flange at 12:00. The engine accelerometers use piezoelectric crystals to sense and transmit radial engine vibration information to the AVM signal conditioner. The piezoelectric crystals are stacked with an inertial mass. When the engine vibrates, the inertial mass tends to stay at rest causing the crystals to be alternately squeezed and released. This produces an electric charge in proportion to the vibration. Metallic collectors receive the electric charges and send them to the AVM signal conditioner. The accelerometers and leads are shielded to prevent interference.

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ENGINE INDICATING SYSTEM

Figure 22 SCL

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AVM System (ENDEVCO) Page: 45

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ALTERNATE NO. 1 BEARING ACCELEROMETER There is an external pad on fan frame strut at the 7:00 position next to the No. 1 bearing accelerometer electrical connector. The pad is used to install an external accelerometer if the internal No. 1 bearing accelerometer fails. This lets vibration monitoring continue until the next scheduled overhaul of the engine.

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ENGINE INDICATING SYSTEM

Figure 23 SCL

JGB

May -- 2001

Nr 1 Accelerometer and Alternal Pad Page: 47

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ENDEVCO AVM SYSTEM OPERATION General The AVM signal conditioner is in the E2--4 rack in the main equipment center. The AVM signal conditioner gets accelerometer and tachometer signals from each engine. Vibration data is sent to EICAS for indication.

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Operation The AVM signal conditioner gets power from the left 115 V ac bus. The piezoelectric accelerometers on each engine generate electrical signals proportional to engine vibration. The AVM signal conditioner gets two accelerometer signals, an N1 speed signal, and an N2 speed signal from each engine. The No. 1 bearing accelerometer signal goes through a tracking filter which passes only vibration signals that are at the same frequency as N1. These vibrations are generated by the N1 rotor near the fan and LPC. The CRF accelerometer signal goes through N1 and N2 tracking filters. Vibrations that match N1 are generated by the LPT. Vibrations that match N2 are generated by the N2 rotor. Both accelerometer signals go through a broad band filter to find the maximum overall vibration level for each engine. The vibration signals are sampled in the multiplexer and sent to the digital signal processing unit. Software then compares the three vibration signals (Fan, LPT and N2) to determine the maximum vibration level for each engine to be shown on the EICAS secondary engine parameter display. If a tachometer signal is not sensed, the broad band vibration signal (BB) appears. The broad band vibration signal also appears when the engine power is below minimum idle. The AVM signal conditioner sends all calculated vibration signals to EICAS for display on the PERF / APU page.

Test The AVM signal conditioner continuously monitors its LRUS. LRU and / or wiring faults are stored in a nonvolatile memory. It there is a fault, a latched message (ENG VIB BITE) appears on the ECS / MSG page, and zeros appear for the vibration indication for the affected engine. The LED on the front of the unit is also turned on. If the fault clears, the vibration indications return. The AVM signal conditioner is tested using the TEST pushbutton switch on the face of the unit. If the test is successful, the red LED on the face of the unit comes on momentarily, then goes out. If the self--test fails, the LED stays on. Faults found during the test are stored in a nonvolatile memory. Monitor faults stored during normal operations, and during self tests, are read using an ARINC reader connected to the AVM signal conditioner front--face connector.

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ENGINE INDICATING SYSTEM

Figure 24 SCL

JGB

May -- 2001

AVM Signal Conditioner Schematic Page: 49

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AVM SYSTEM INDICATIONS General Engine vibration data appears on the EICAS secondary engine display directly below the oil quantity indications. The indications consist of a vibration mode callout, and the vibration value using both a digital readout and a vertical analog pointer. The vibration data also appears on the PERF/APU page. Vibration Mode A white FAN, LPT, N2 or BB callout appears above the actual readout to identify the source of the highest vibration. Vibration Data A digital indication of engine vibration appears as a white number enclosed in a white box next to the vertical scale. The readout indicates engine vibration in the unitless range O to 5. A white triangular pointer on the inside of a vertical scale also indicates engine vibration level. There are two digital and vertical scale indications, one for each engine.

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PERF / APU Page The FAN, LPT, N2 and BB vibration levels are all shown on the PERF / APU page. Airplanes with EICAS computers with part number S 242N701704 and later also display the vibration phase angle for FAN and LPT vibrations. The phase angle is used for engine trim balance computations. See MM 72--31--00/501.

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ENGINE INDICATING SYSTEM

Figure 25 SCL

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AVM System Indications Page: 51

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NOTES :

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ATA 80 / 74 START & IGNITION TABLE OF CONTENT General Description Start System Components Start Air Sources Starter Start Valve Start Supply Duct Starting System Opeation Ignition Exciter Ignition Leads Ignition Plugs Ignition System Power Ignition System Control Engine Motoring

002 004 008 010 012 014 016 018 020 022 024 026 028

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ENGINE STARTING AND IGNITION General Description The Start and Ignition System are normally working together, because start system is interlock for the ignition system. Starting The engine starting system turns the N2 rotor to start the engine. The N2 rotor is turned by the pneumatic starter through the horizontal and radial drive shafts. The system can be used in the air or on the ground. The starting system is also used to motor an engine on the ground. Pneumatic power is, from any of three sources: -- Pneumatic ground carts (2 connectors) -- Auxiliary Power Unit (APU) -- Cross bleed air from an operating engine. System components for each engine include the pneumatic starter and starter control valve. Switches on the ignition and start control panel control operation of the engine starting system.

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Ignition The ignition system supplies the high energy spark to start or sustain combustion of the fuel / air mixture in the combustor. Each engine ignition system has two electrically and physically independent circuits. Each circuit has an ignition exciter connected to an igniter plug by a shielded lead.

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ENGINE START & IGNITION SYSTEMS

Figure 1 SCL

JGB

May -- 2001

Engine Starting and Ignition Page: 3

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ENGINE START & IGNITION SYSTEMS

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STARTING AND IGNITION SYSTEMS General The starting system turns the engine to reach the rotor speed necessary to initiate self--sustained engine operation. The ignition system ignites the fuel / air mixture in the combustor during starting, and helps sustain ignition during selected low--power operations. The starting system includes a pneumatic starter control valve and a pneumatic starter. A VALVE light on the engine ignition and start control panel indicates a disagreement between the commanded position and actual position of the pneumatic starter control valve. The light comes on momentarily when the valve is in transit. The ignition system has two ignition exciters (1 and 2) and two igniter plugs (1 and 2). Operation The EEC active channel controls starting and ignition in response to control switch input from the engine ignition and start control panel on the P5 overhead panel, and from the fuel control switches on the P10 panel. The ignition select switch allows a choice of using a single igniter plug or both igniter plugs for both engines. The ignition / start control switches control the pneumatic starter control valve and allow four additional choices of ignition use. When pneumatic power is available, the starter is powered by moving the ignition / start control switch to GND. This also enables ignition. The fuel control switch is normally moved to RUN at 20 % (15 % minimum) of N2. This permits fuel flow to the combustor, and turns on ignition. At 50 % of N2, the ignition / start control switch automatically moves to AUTO. This closes the pneumatic starter control valve and normally turns off ignition.

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ENGINE START & IGNITION SYSTEMS

Figure 2 SCL

JGB

May -- 2001

Starting and Ignition Systems Schematic Page: 5

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STARTING SYSTEM COMPONENT LOCATIONS Pneumatic Starter The pneumatic starter is mounted on the aft side of the accessory gearbox in the 6:00 position. It turns the N2 rotor to start the engine. It has ports for servicing and for a magnetic chip detector. Pneumatic Starter Control Valve The pneumatic starter control valve is mounted aft of the pneumatic starter between the starter inlet and the air supply duct. The valve controls the flow of air to the starter. A filter protects the valve actuator.

For Training Purposes Only

Engine Ignition and Start Control Panel The engine ignition and start control panel is on the P5 overhead panel. The panel includes the ignition / start control switches, the ignition select switch, and two in--transit or disagreement valve lights.

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ENGINE START & IGNITION SYSTEMS

Figure 3 SCL

JGB

May -- 2001

Starting System Component Locations Page: 7

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ENGINE STARTING SYSTEM -- AIR SOURCES The pneumatic sources available for starting an engine induce the APU, ground air sources, or the opposite operating engine. The nominal required pressure for starting an engine is 45 psig. A pneumatic control panel on the P5 overhead panel includes the switches and indications necessary to control and monitor the air source selection. For normal starting, the isolation valve and APU air supply valve switches are all latched in. Valve operation during starting is then automatic. The left (right) engine PRSOV switch is only latched in if the engine is running and supplying the air supply to start the other engine.

Starter Control Valve Malfunctions If the engine pneumatic starter control valve fails to close after an engine start, the air source to the pneumatic starter can be removed by manually closing the proper isolation valves or engine PRSOV. If a ground air source is in use and the left pneumatic starter control valve fails to close, it is necessary to disconnect the ground air source to stop airflow to the pneumatic starter.

Ground Air Sources There are two ground pneumatic service connections. Access is through a left forward wing--to--body fairing door. The rotary switch that controls the left engine pneumatic starter control valve must be turned to GND to enable automatic left engine starting. The rotary switch that controls the right engine pneumatic starter control valve must be turned to GND, and the right isolation valve switch must be latched in to enable automatic right engine starting.

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APU Air Source The APU supplies air at 40 to 50 psi for engine starting. During an engine start, the APU operates at a higher speed (101%) to supply additional air flow. The switches that control the APU air supply valve and center isolation valve must be latched in to enable automatic valve operation.The ignition / start control switch for the engine being started must be turned to GND to begin the start process. Operating Engine Air Source When an operating engine is used as a pneumatic source to start the other engine, 8th stage bleed air is used at high power settings (N2 greater than 75 %) and l4th stage bleed air is used at low power settings (idle to 75 % of N2). The high pressure valve automatically selects 8th or l4th stage air. Air pressure is regulated by the pressure regulating valve.

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ENGINE START & IGNITION SYSTEMS

Figure 4 SCL

JGB

May -- 2001

Engine Starting System - Air Sources Page: 9

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PNEUMATIC STARTER (HAMILTON STANDARD) General The pneumatic starter is a single stage, axial flow, turbine air motor mounted with a V--band clamp to the the aft side of the accessory gearbox between the fuel pump and the IDG. Two locator pins are used to align the starter. The starter weighs about 35 pounds (16 Kg). A filter plug is located on each side of the starter. A pressure fill fitting, overflow plug, and drain a plug with a magnetic chip detector are on the bottom.

Oil Servicing To add oil, remove the overfill plug and pour oil through the oil filler plug port, or pump; oil through the pressure fill fitting, until it flows from the overfill port. Magnetic Chip Detector The magnetic drain plug assembly has an inner magnetic probe and an outer drain plug. A check valve in the plug permits the magnetic probe to be removed for inspection without draining the oil.

Starter Operation When pneumatic power is available at the starter inlet, the turbine turns the N2 rotor through the gear train, clutch, spline drive and gearbox. The clutch allows the starter to coast to a stop when pneumatic power is shut off. When N2 is greater than about 40 %, centrifugal force holds the clutch pawls away from the turbine drive teeth. Below this speed the pawls ratchet against the teeth. The starter may be engaged normally when below N2 is 20 %, and in case of fire, when N2 is below 30 %. CAUTION:

STARTER RE--ENGAGEMENT ABOVE 30 % OF N2 CAN RESULT IN STARTER OR GEARBOX DAMAGE.

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Duty Cycle Limitations The normal duty cycle is 1 minute on and 30 seconds off. The extended duty cycles are as follows: -- 0 -- 5 minutes on -- disengage starter and permit N2 to go to zero before re--engagement. -- 5 -- 10 minutes on -- follow with a 20 minute starter cooling period. -- 10 -- 15 minutes on -- follow with a 30 minute starter cooling period. Removal and Installation Support the starter during removal and installation, to avoid damage to the gearbox or starter. CAUTION:

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TO AVOID EXCESSIVE LOADS ON INTERNAL PARTS, DO NOT LIFT STARTER BY DRIVE SHAFT.

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ENGINE START & IGNITION SYSTEMS

Figure 5 SCL

JGB

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Pneumatic Starter (Hamilton Standard) Page: 11

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STARTER CONTROL VALVE (HAMILTON STANDARD) General The starter control valve is a spring--loaded closed butterfly--type. It is solenoid controlled and pneumatically powered. It can be manually operated using a square drive tool. There are valve open and valve closed position switches. The valve is mounted on the starter inlet. The starter air supply duct is connected to the valve inlet.

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Operation With pneumatic power available, the EEC energizes the solenoid to open the starter control valve. The valve sends position feedback to the EEC through the position switches. For manual operation the valve is reached with a 3/8 inch square drive through a hole in the thrust reverser latch access door. Instructions are on the door. The valve must be held open against a spring. Communication with the flight compartment must be maintained all time. WARNING:

WHEN MANUALLY OPERATING CONTROL VALVE, WEAR HAND AND ARM COVERS. HEAT AND AIR BLAST EXHAUST FROM STARTER COULD INJURE PERSONNEL.

CAUTION:

STARTER MAY BE DAMAGED IF VALVE IS NOT CLOSED WHEN N2 IS GREATER THAN 50 %.

CAUTION:

MANUAL OPERATION OF STARTER CONTROL VALVE WITHOUT PNEUMATIC PRESSURE IN THE DUCT MAY DAMAGE VALVE.

Maintenance Practices An air filter on the valve is cleanable. A dirty filter results in slowor sluggish valve opening. The filter element is located behind a filter cap, packing, and spring.

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ENGINE START & IGNITION SYSTEMS

Figure 6 SCL

JGB

May -- 2001

Starter Control Valve (Hamilton Standard) Page: 13

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PNEUMATIC STARTER SUPPLY DUCT General The pneumatic starter supply duct directs air from the airplane pneumatic manifold to the start control valve. The duct consists of two sections that are coupled together and mounted to the left side of the compressor rear frame with support links. Removal and Installation To remove the starter supply duct, open the left thrust reverser half. Remove the V--band clamps and pressure seals to remove the upper supply duct from the pneumatic interface duct and the lower pneumatic supply duct. Remove the V--band clamps and seals, and the upper and lower support links, to remove the lower supply duct from the upper supply duct and the start control valve. CARE MUST BE TAKEN NOT TO DAMAGE THE SEALS WHEN CONNECTING THE STARTER DUCT COUPLINGS.

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CAUTION:

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ENGINE START & IGNITION SYSTEMS

Figure 7 SCL

JGB

May -- 2001

Pneumatic Starter Supply Duct Page: 15

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ENGINE START & IGNITION SYSTEMS

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ENGINE STARTING SYSTEM OPERATION Operation The engine ignition / start control switch is placed to GND to begin the engine start sequence. A holding coil in the switch is energized to keep the switch in GND if the speed card senses N2 less than 50 %. The switch is released and snaps back to AUTO when the coil is de--energized. The switch may be released from GND manually if necessary. When the switch is in GND, ENGINE START 1 is also energized if N2 is less than 50 %. This energizes ENGINE START 3, causing the enabled channel of the EEC to energize the pneumatic starter control valve solenoid. The valve opens, allowing pneumatics to the starter. Indications Either of the following conditions causes the VALVE light on the engine ignition and start control panel to come on: The engine starting system commands the pneumatic starter control valve to open (ENG START 1 energized), but the valve is not fully open. This occurs if the valve fails, or at the beginning of the start sequence while the valve opens. The engine starting system commands the pneumatic starter control valve to close (ENG START 1 relaxed), but the valve is not fully closed. This occurs if the valve fails, or at the end of the start sequence (50 % of N2) while the valve closes. The engine starting system continues to command the pneumatic starter control valve to open (ENG START 1 stays energized), and the valve stays fully open, for 2 seconds after N2 has reached 52 %. If the pneumatic starter control valve does not open fully within 5 seconds after the ignition/start control switch is moved to GND, the EICAS level C message L (R) ENG STARTER appears. This also causes the VALVE light to stay on.

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The EICAS level B message L (R) STARTER CUTOUT appears 5 seconds after either of the following: -- The engine starting system commands the pneumatic starter control valve to close (ENG START 1 relaxed), but the valve does not fully close. -- N2 reaches 52 % and the engine starting system continues to command the pneumatic starter control valve to open (ENG START 1 stays energized). The EICAS level B message L (R) STARTER CUTOUT removes all other current EICAS level B and C messages, and inhibits new level B and C messages for 20 seconds. If the L (R) STARTER CUTOUT message appears, close the proper isolation valves to remove pneumatic power from the starter. If the L STARTER CUTOUT message appears while starting the left engine using ground air sources, the ground air source must be removed. CAUTION:

IF VALVES IS NOT CLOSED WHEN N2 IS GREATER THAN 50 % RPM, STARTER MAY BE DAMAGED.

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ENGINE START & IGNITION SYSTEMS

Figure 8 SCL

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May -- 2001

Engine Starting System Operation Page: 17

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IGNITION EXCITERS The two identical ignition exciters convert 115 V , 400 Hz ac power to a 14--to--18 kilovolt pulsed output at the rate of approximately one pulse per second. The exciters normally get power from the main ac buses. Alternatively, the exciters can receive power from the standby ac bus. The EEC controls the source of power for the ignition exciters. The exciters are rated for continuous operation. Each exciter is a hermetically sealed unit with two connectors. One connector receives power froin the EEC. The other connector sends power to the igniter through the ignition lead. The exciters are below the EEC (not shown) on the lower left side of the fan case. Access is through the fan cowl. Exciter No. 1 is above exciter No. 2. Exciter No. 1 powers igniter plug 1, and exciter No. 2 powers igniter plug 2. IGNITION VOLTAGE IS DANGEROUSLY HIGH. TOUCHING ELECTRICAL CONTACTS MAY BE FATAL. IGNITION MUST BE 0FF FOR SEVERAL MINUTES AND EXCITER GROUNDED BEFORE TAKING OUT IGNITION COMPONENTS.

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WARNING:

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ENGINE START & IGNITION SYSTEMS

Figure 9 SCL

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Ignition Exciters Page: 19

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IGNITION LEADS The ignition leads carry electrical power from the ignition exciters to the igniter plugs. Both leads go from the ignition exciters, through the pylon fire seal, to the igniter plugs. Access to the ignition leads is through the fan cowls and the right thrust reverser half. The conductor is 14 AWG stranded copper with silicone rubber insulation within a flexible conduit. The conduit has an inner copper braid and an outer nickel braid. There is a plastic sleeve over the cold section of the lead and an air cooling jacket / conduit over the hot section. Fan air from the turbine case cooling duct cools the leads. After cooling the lead, the air goes through a port just above the coupling nut to cool the igniter plug. Observe safety precautions when removing or handling the ignition leads. High voltage can be present.

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ENGINE START & IGNITION SYSTEMS

Figure 10 SCL

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Ignition Leads Page: 21

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IGNITER PLUGS The igniter plug is a surface gap type plug. The No. 1 igniter plug is at the 5:00 position. The No. 2 igniter plug is at the 3:30 position. Access to the plugs is through the right thrust reverser half. The plugs are threaded into adapters bolted to the CRF; factory--installed gasket spacers between the adapter and CRF ensure proper plug depth. A clamped igniter shroud protects and cools the igniter plug. When igniter plugs are removed, the plug can be replaced or reinstalled without installing gasket spacers1 as long as the adapter is not removed. There is an integral gasket which must be installed on the plug prior to installation into the adapter. Refer to M.M. 74--21--02 for removal and installation. IGNITION SYSTEM VOLTAGE IS DANGEROUSLY HIGH. IGNITION / START CONTROL SWITCH MUST BE 0FF BEFORE REMOVING ANY IGNITION COMPONENTS. ALLOW SEVERAL MINUTES TO ELAPSE BETWEEN OPERATION OF IGNITION SYSTEM AND REMOVAL OF COMPONENTS. WHEN DETACHING CABLE FROM IGNITER PLUGS, DISCHARGE CURRENT BY GROUNDING CABLE TERMINAL TO ENSURE COMPLETE DISSIPATION OF ENERGY FROM THE SYSTEM. FAILURE TO FOLLOW THIS PROCEDURE COULD RESULT IN SEVERE INJURY TO PERSONNEL.

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WARNING:

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ENGINE START & IGNITION SYSTEMS

Figure 11 SCL

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Igniter Plugs Page: 23

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IGNITION SYSTEM POWER Power Sources When the engine is running, the left and right 115 volt ac buses normally supply ignition system power to both channels of the EEC. If a bus power sense relay is relaxed, the standby bus supplies ignition system power to the EEC. Power is supplied to the EEC when the fuel control switch is in RUN and the engine fire switch is in the NORMAL position. The EEC supplies power from the left bus to one ignition exciter, and supplies power from the right bus to the other exciter. This ensures that one ignition exciter on each engine can operate if one of the main buses does not have power. Power Control The EEC controls the power to the ignition exciters based on which EEC channel is active, switch settings on the engine ignition and start control panel, and N2 speed. The ignition selection logic within each EEC channel alternates between the two ignition exciters on each engine start to ensure even wear of the igniters.

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Indications If the standby bus is supplying power for ignition, the EICAS maintenance message IGN 1 (2) STBY BUS appears on the ECS / MSG page. If IGN 1 STBY BUS appears and the upper EICAS display unit is operational, the left ac bus has power but the power sense circuits have malfunctioned. If the left ac bus does not have power, the upper EICAS display unit is blank and IGN 1 STBY BUS appears on the lower EICAS display unit. The IGN 2 STBY BUS message and the lower EICAS display unit operate the same way.

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ENGINE START & IGNITION SYSTEMS

Figure 12 SCL

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May -- 2001

Ignition System Power Page: 25

Lufthansa LAN Technical Training For Training Purposes Only

ENGINE START & IGNITION SYSTEMS

BOEING -- 767 / 300 CF6 -- 80C2 80 / 74 -- 00

IGNITION SYSTEM CONTROL General The EEC supplies 115 V ac power to one or both ignition exciters based upon ignition system commands. The commands to EEC are ground signals that enable SINGLE or DUAL ignition. If the EEC does not sense a SINGLE or DUAL ground signal1 neither ignition exciter gets power. The ignition enabling command is controlled by the ignition select switch, ignition / start control switch, engine fire switch, fuel control switch, engine thermal anti--ice relay, and flap position proximity switch. Operation Ignition is only enabled when the two fuel/ignition control relays are relaxed. This occurs when the engine fire switch on the P10 panel is in the NORMAL position, and the fuel control switch on the P10 panel is in RUN. Dual ignition is enabled when the ignition select switch on the P5 overhead panel is in the BOTH position, or when the ignition / start control switch on the P5 panel is in the FLT position. Single ignition is enabled in all other cases that ignition is commanded. When the engine fire switch is in the NORMAL position, and the fuel control switch is in RUN, ignition is enabled based on the position of the ignition / start control switch: -- GND: ignition is enabled, and the pneumatic starter control valve is opened until N2 reaches 50 %, when the switch automatically moves to the AUTO position. -- AUTO: ignition is enabled when the engine thermal anti--ice system is on (bad weather), or when the flaps are down (takeoff and landing). -- 0FF: ignition is disabled. -- CONT: ignition is continuously enabled. -- FLT: ignition (DUAL with SINGLE as backup) is enabled for in--flight starts.

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Page: 26

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ENGINE START & IGNITION SYSTEMS

Figure 13 SCL

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May -- 2001

Ignition System Control Page: 27

Lufthansa LAN Technical Training

ENGINE START & IGNITION SYSTEMS ENGINE MOTORING General The engine motoring procedure is used for any operation that requires engine rotation. Use dry motoring for all tests that require engine motoring unless wet motoring is specifically required. Dry Motoring Before dry motoring the engine, perform premotoring procedures listed in the maintenance manual. -- On the main power circuit breaker panels, open the appropriate circuit breakers. -- Make sure the fuel control switches are in CUTOFF, and the thrust levers are at IDLE. Check for full engine EICAS displays on upper and lower display units. -- Turn the engine start switch to GND. The start switch is electrically latched to the GND position. The switch is automatically released at 50 % of N2, or it can be manually released prior to reaching 50 % of N2. Maximum motoring speed is 30 to 34 % of N2. OBSERVE STARTER LIMITATIONS PER MAINTENANCE MANUAL. Confirm indication of N1 rotation and oil pressure on EICAS display. Turn the start switch to 0FF. CAUTION:

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BOEING -- 767 / 300 CF6 -- 80C2 80 / 74 -- 00

-- Check that the fuel control switches are in CUTOFF, and the thrust levers are at IDLE. Check for full engine EICAS displays on upper and lower display units. -- Turn the forward and aft fuel boost pump switches ON. -- Turn the start switch to GND. The start switch is electrically latched to the GND position. The switch automatically releases at 50 % of N2, or it can be manually released prior to reaching 50 % of N2. Maximum motoring speed is 30 to 34 % of N2. OBSERVE STARTER LIMITATIONS PER MAINTENANCE MANUAL. At 15 % of N2, move the fuel control switch to RUN and continue to wet motor until 550 PPH (250 KGPH) indicated fuel flow occurs or for 60 seconds. Fuel fogging from the engine gas path should occur. Confirm indication of N1 rotation and oil pressure on EICAS display. Move the fuel control switch to CUTOFF and continue to motor for 30 additional seconds (minimum) or until vapor ceases at engine exhaust. Turn the start switch to 0FF.

CAUTION: --

----

Wet Motoring Wet motoring lets fuel into the combustion chamber. -- Before wet motoring the engine, perform premotoring procedures listed in the maintenance manual. -- On the main power circuit breaker panel, open the appropriate circuit breakers. CAUTION:

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DO NOT LEAVE IGNITION CIRCUIT BREAKERS CLOSED. INADVERTENT LIGHT UP COULD OCCUR.

May -- 2001

Page: 28

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BOEING -- 767 / 300 CF6 -- 80C2 80 / 74 -- 00

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ENGINE START & IGNITION SYSTEMS

Figure 14 SCL

JGB

May -- 2001

Motoring Control and Indications Page: 29

Lufthansa LAN Technical Training

ENGINE START & IGNITION SYSTEMS

BOEING -- 767 / 300 CF6 -- 80C2 80 / 74 -- 00

IN FLIGHT START In--FIight Start Data When an engine is shut down in flight, the in -- flight start envelope appears in the Iower left portion of the screen. The envelope shows an airspeed range for the current flight level and the next two flight levels below. The highest flight level that can be shown in FL 30.0 (30,000 feet). A fuel on command bug ( index ) appears on the Iower display N2 gage and a cross--bleed capability is available in the air.

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In -- Flight Start Display There are four conditions to satisfy for the in -- flight start display to appear: -- Primary and secondary engine data is displayed. ( It secondary data is not displayed and a fuel control switch is moved to CUTOFF while in the air, secondary parameters are automatically commanded.) -- The airplane is in the air ensured by both system 1 and 2 air / ground relays. -- Either fuel control switch is in CUTOFF. -- The aftected engine fire switch not pulled. There are four conditions that can cause removal of the in -- flight start display: -- The airplane is on the ground. -- The affected engine fire switch pulled. -- The engine is running. -- Primary and secondary engine data is not displayed. Cross--Bleed Message Display When the in -- flight start envelope is showed on the upper DU and the airpIane is within that envelope the message “X--BLD” in magenta appears in the upper right corner of the Iower DU if engine speed minimums are not met. The message will be automatically removed by the same Iogic that controls the in -- flight start display.

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Page: 30

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BOEING -- 767 / 300 CF6 -- 80C2 80 / 74 -- 00

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ENGINE START & IGNITION SYSTEMS

Figure 15 SCL

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May -- 2001

In - Flight Start Data Page: 31

BOEING -- 767 / 300 CF6 -- 80C2 80 / 74 -- 00

NOTES :

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ENGINE START & IGNITION SYSTEMS

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May -- 2001

Page: 32

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

ATA -- 78 ENGINE EXHAUST TABLE OF CONTENT General Description Translating Cowls Track Liners and Sliders Blocker Doors and Drag Links Deflectors PRSOV DPV and Pressure Switch CDU Rotary Flexible Drive Shaft Angle Gearbox and Ballscrew Actuators Electromechanical Brake CDU Feedback Transducer CDU Position Switches Thrust Lever Interlock Actuator Thrust Reverser Control Switches Thrust Reverser Operation Thrust Reverser Electrical Operation Thrust Reverser Indicating System Translating Cowl Manual Operation Translating Cowl External Air Powered Translating Cowl PRSOV Hold--Open Translating Cowl Ground Service Switch T/R Deactivation Lockout T/R Cowls Opening System T/R Power Pack and Control Switches T/R Cowls System Operation

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002 006 008 010 012 014 016 018 022 024 026 032 034 036 040 042 046 048 050 052 054 056 058 060 062 064

Page: 1

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER SYSTEM General The thrust reverser, when deployed, redirects fan air forward to decelerate the airplane. The thrust reverser is normally deployed during landing rollout or during a rejected takeoff. Each engine has two thrust reverser halves. Each half includes a translating cowl, six blocker doors with drag links, 16 deflectors, and a center drive unit (CDU) with three actuators, two of which are driven through flexible drive shafts and angle gearboxes. The two translating cowls operate independently. When the thrust reverser is stowed, the translating cowl fairs with the fan cowl and the blocker doors are retracted. In the stowed position, the thrust reverser directs fan air aft for forward thrust. When the thrust reverser is deployed, the translating cowl slides aft to expose the deflectors and to block the fan air path with the blocker doors. This directs fan air forward, reversing the direction of thrust. Turbine exhaust air is not reversed. While the fan air is deflected forward to provide deceleration, turbine exhaust is still providing some forward thrust.

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ENGINE EXHAUST

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May -- 2001

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ENGINE EXHAUST

Figure 1 SCL

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May -- 2001

Thrust Reverser System Page: 3

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER SYSTEM OPERATION Deploy When reverser deployment is commanded, switch and relay logic provide power to unlock the electro--mechanical brake and to open the T/R PRSOV. Air from the T/R PRSOV flows to the left and right CDU‘s and to the DPV. An air signal from the DPV to the CDU arms the CDU to the deploy mode. Air motors in the CDU‘s drive ballscrew actuators attached to the center of the translating cowls. Angle gearbox and ballscrew actuators are attached to the upper and lower ends of the translating cowls. Flexible drive shafts mechanically connect the angle gearbox and ballscrew actuators to the CDU‘s. The air motors in the CDU‘s drive the center ballscrew actuators and the upper and lower flexible drive shafts. The flexible drive shafts then drive the upper and lower angle gearbox and ballscrew actuators. The ballscrews move the translating cowls aft. Blocker doors, pulled by the drag links, rotate from a flush position against the inside of the translating cowl to a position blocking the fan air discharge path. The fan air discharge is redirected forward through the deflectors. The air motors in each CDU also drive a CDU position feedback transducer for thrust reverser position feedback to the EEC. The EEC then sends a signal to the thrust reverser interlock actuator to permit increased reverse thrust. Stow When the thrust reverser is commanded to stow, air from the T/R PRSOV flows to the left and right CDU‘s and the DPV. Now the DPV remains closed, blocking the air signal to the CDU‘s. This arms the CDU‘s to the stow mode. The air motors reverse direction, driving the actuators and translating cowl forward to the stow position. The blocker doors (pushed by the drag links) rotate back to a flush position with the inner translating cowl. When fully stowed, the system deenergizes the solenoids on the electro--mechanical brake. The system is now locked in the stowed position by the CDU cone brakes and by the electro mechanical brakes.

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May -- 2001

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ENGINE EXHAUST

Figure 2 SCL

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May -- 2001

Thrust Reverser System Operation Page: 5

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

TRANSLATING COWL General When the thrust reverser is stowed, the translating cowl covers the deflectors and acts as a section of power plant cowling. When the reverser is deployed, the translating cowl slides aft (translates) to uncover the deflector segments. The translating cowl is constructed of a Kevlar, graphite, and fiberglass facesheet with a Nomex core. Hinges are bonded into the inner wall. There are six blocker doors attached to the hinges in the inner wall of each cowl. Reverser Track Fairing The thrust reverser track fairing permits smooth airflow over the thrust reverser sliders and liners. The fairing is on the top and bottom of each translating cowl. Maintenance Practices If the translating cowl needs to be removed it must first be deployed about 6--8 inches. Remove the actuator access panels. DO NOT REMOVE CLEVIS PIN RETAINING CLIP BOLT. BACK BOLT OUT ENOUGH TO ROTATE RETAINING CLIP. REMOVAL OF BOLT WILL DAMAGE NUTPLATE. Loosen the retaining clip bolt and turn the clip. Remove the clevis pins to disconnect the actuators from the translating cowl. CAUTION:

DO NOT OPEN THRUST REVERSER HALF BEYOND THE 34 DEGREE POSITION WITH THE TRANSLATING COWL EXTENDED. DAMAGE TO TRANSLATING COWL OR STRUT MAY RESULT. Open the thrust reverser half to the 20° position. Disconnect the blocker door drag links from the aft side of the blocker doors. Slide the translating cowl aft on its track.

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CAUTION:

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

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ENGINE EXHAUST

Figure 3 SCL

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May -- 2001

Translating Cowl Page: 7

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

TRACK LINERS AND SLIDERS General The translating cowl slides on a low friction track and slider mechanism. There are two sets of thrust reverser track sliders on each translating cowl, one on the top of the reverser half and one on the bottom. Each set has a T-shaped main slider and a J--shaped auxiliary slider. These sliders move on T-and J--shaped track liners. The liners are fixed to the stationary fan duct that holds the translating cowl to the thrust reverser. The J--shaped auxiliary track liner is mounted above the deflectors. The T--shaped main track liner is mounted under the deflectors.

For Training Purposes Only

Maintenance Practices Thrust reverser track sliders must be inspected periodically for wear. The low friction Teflon surfaces must be smooth and clean.

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May -- 2001

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ENGINE EXHAUST

Figure 4 SCL

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May -- 2001

Track Liners and Sliders Page: 9

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

BLOCKER DOORS AND DRAG LINKS General Each thrust reverser half has six blocker doors mounted to the inner wall of the translating cowl. The blocker doors deflect fan air radially outward when the translating cowl is deployed. The drag links pull the doors into position during deployment. The doors are made of fiberglass and graphite composite, with bonded aluminum hinges. There are two hinges on the wide, forward end that connects to the inner wall of the translating cowl. There is a drag link connection in the center of the door. The drag link is pinned to this connection and is spring--loaded to hold the door closed when the reverser is stowed. All six blocker doors are interchangeable.

The drag link can also be removed at this time by cutting away the protective coating from the nuts and washers on the inboard side of the inner fan duct cowl and removing the nuts and washers. If necessary, cut the coating around the link support. Pull the link support and drag link out of the fan duct cowl. Separate the drag link from the link support by removing the bolt, washer and link pin.

Maintenance Practices The blocker doors must be checked for movement at the point of attachment. If removal is necessary, manually deploy the translating cowl about 16 inches. DO NOT OPEN THE THRUST REVERSER HALF BEYOND THE 23° POSITION WHEN THE THRUST REVERSER TRANSLATING COWL IS DEPLOYED. DAMAGE TO TRANSLATING COWL OR STRUT MAY RESULT. Open the thrust reverser half to the 23° position.

CAUTION:

RELIEVE SPRING PRESSURE BY ALTERNATELY LOOSENING SPRING RETAINER CLIP SCREWS. REMOVING ONE SCREW BEFORE LOOSENING THE OTHER COULD RESULT IN INJURY TO PERSONNEL FROM SPRING RELEASE UNDER PRESSURE. Release spring pressure by alternately loosening the two spring retainer clip screws. Remove the clip and springs. Disconnect the drag link from the blocker door by pushing the blocker door forward over the drag link and removing the bolt. Remove the bolts that attach the hinges to the translating cowl and remove the blocker door.

For Training Purposes Only

WARNING:

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May -- 2001

Page: 10

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

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ENGINE EXHAUST

Figure 5 SCL

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May -- 2001

Blocker Doors and Drag Links Page: 11

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

DEFLECTORS General There are 16 deflectors on each thrust reverser half that direct fan air forward when the thrust reverser is deployed. When the reverser is stowed, the translating cowls cover the deflectors. When the reverser is deployed, the blocker doors direct fan air through the deflectors. The deflectors are made of cast aluminuin. The front and rear edges of the deflectors are bolted to the thrust reverser fixed structure. There are gang channels between the deflectors to interconnect the deflectors. The gang channels are screwed to the deflectors with tri--wing screws. The top deflector has two gang channels. Five different types of deflectors are mounted on each thrust reverser half. Each type directs the air differently as shown. Deflectors are also called cascade segments or cascade vane segments. Maintenance Practices Thrust reverser deflectors are not interchangeable because of the different flow angles. Exact deflector position is found in the maintenance manual. Deflectors must be inspected periodically for cracks, corrosion, and impact damage. DO NOT OPERATE ENGINE IN REVERSE THRUST WITH DEFLECTORS MISSING. DAMAGE TO THE REVERSER MAY RESULT.

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CAUTION:

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May -- 2001

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

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ENGINE EXHAUST

Figure 6 SCL

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May -- 2001

Deflectors Page: 13

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER PRESSURE REGULATING AND SHUTOFF VALVE The thrust reverser (T/R) pressure regulating and shutoff valve (PRSOV) isolates the thrust reverser pneumatic system from the airplane pneumatic system, and regulates the pressure. There is one valve in each strut at the entrance to the reverser supply duct downstream of the precooler. Access is through a pressure relief door on the right side of the strut. The T/R PRSOV has a steel valve body with a poppet valve, a solenoid valve, a pressure regulator, and a relief valve. The poppet valve is spring--loaded closed. When reverse thrust is selected, the solenoid valve is energized. Air flows around the poppet valve stem, through the solenoid valve, and pressurizes the pneumatic actuator. This opens the poppet valve. The pressure regulator opens when the inlet pressure is higher than 70 psig. This modulate the poppet valve, regulating downstream pressure. Normally, the air supply pressure is not high enough to require valve regulation. However, the engine may develop enough 8th stage bleed pressure to open the regulator during a rejected takeoff. The relief valve opens if actuator pressure exceeds 150 psig.

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ENGINE EXHAUST

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May -- 2001

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ENGINE EXHAUST

Figure 7 SCL

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May -- 2001

T/R PRSOV Page: 15

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

DIRECTIONAL PILOT VALVE AND PRESSURE SWITCH General The directional pilot valve (DPV) changes the direction of the directional control valve (DCV) in the CDU to control thrust reverser deploy and stow. The DPV pressure switch completes a circuit for thrust reverser indication. The DPV and pressure switch are on the torque box of the left reverser half. There is one per engine. Access is through the left fan cowl panel. The DPV is spring--loaded closed. It has a ball and poppet valve on a common shaft, a solenoid, and a cleanable air filter. The pressure switch is a two position microswitch.

For Training Purposes Only

Operation The DPV either pressurizes or vents the directional control valve actuator inside both CDU‘s for an engine. When the solenoid is de--energized, air pressure from the T/R PRSOV is blocked and air from the directional control valve is vented throught the DPV ball valve to ambient. When reverse thrust is selected, the solenoid is energized and the ball valve moves down, closing the vent. The poppet valve opens, permitting air pressure from the T/R PRSOV to go to the directional control valve. The pressure switch senses air pressure to the DPV. It is open when the T/R PRSOV is closed. The pressure switch closes when it senses pressure from the T/R PRSOV. Its position is independent of the directional pilot valve position. There is an indication in the flight compartment if the pressure switch position disagrees with the T/R PRSOV position. This indication is discussed later.

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May -- 2001

Page: 16

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

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ENGINE EXHAUST

Figure 8 SCL

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May -- 2001

DPV and Pressure Switch Page: 17

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

CENTER DRIVE UNIT General The center drive unit (CDU) is a pneumatic motor with a ballscrew actuator for deploying and stowing the thrust reverser. The CDU has a position switch module, a gearbox and a position feedback transducer. The gearbox has two flexible drive shaft output drives and a manual drive pad. One CDU is mounted on each thrust reverser half between the upper and lower angle gearboxes. Access is through the fan cowl. The CDU‘s require an air supply connected to either end of the inlet tee fitting. The other end is normally capped, but can be used for a ground air supply. The actuator stroke length is 22 inches. The position indicating switch module is line replaceable and does not require rigging. The manual brake release lever releases the cone brake for manual operation of the translating cowl. The brake releases when the lever is moved about 60° into a detent. The fan cowl automatically closes the lever if it is left in the 60° detent position. The gearbox has two splined output drives that turn the flexible drive shafts. It also turns the CDU position feedback transducer and has a square drive pad for manual operation. Removal Remove middle actuator access panel. Manually deploy the thrust reverser half about 6--8 inches until the ballscrew actuator clevis pin is exposed. Deactivate the thrust reverser by reversing the lockout plate. Loosen the retaining clip bolt. Rotate clip and remove clevis pin using a pin extracting tool. DO NOT REMOVE CLEVIS PIN RETAINING CLIP BOLT. BACK BOLT OUT ENOUGH TO ROTATE RETAINING CLIP. REMOVAL OF BOLT WILL DAMAGE NUTPLATE. Disconnect the rotary flexible drive shafts and remove the 4 CDU flange bolts. Ensure that the CDU upper flexible drive shaft does not slide out of the sheath. Pull CDU and ballscrew actuator from torque box noting shim installation details. Mark the position of the actuator on the ballscrew to aid CDU installation.

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CAUTION:

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

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ENGINE EXHAUST

Figure 9 SCL

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May -- 2001

CDU Page: 19

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

CENTER DRIVE UNIT OPERATION General The directional control valve (DCV) includes a directional valve, a helix rod and spring, and a valve actuator piston. The DCV is springloaded in the stow position. The actuator cone brake has a spring--loaded friction cone and rotating mating cone mounted on the air motor shaft. The valve actuator piston moves a pivoted lever to release the brake. When the brake is engaged, the air motor can rotate in the stow direction, but not in the deploy direction. The ballscrew and ballnut actuator is one assembly. The air motor turns the ballscrew. The ballscrew is free to rotate, but can not translate. It engages the ballnut actuator. The ballnut actuator is free to translate but can not rotate because it is attached to the translating cowl. The stop rod is linked to the DCV assembly on one end and has a mushroom shaped head on the other. It turns the DCV through an override linkage, operates the CDU position indicating switch assembly, and keeps the cone brake from engaging until the cowl is completely stowed. The CDU position indicating switch assembly has stow and deploy limit switches to indicate thrust reverser position. The switches also control electrical power to the T/R PRSOV. They are operated by the stop rod.

Stow Operation The air signal from the DPV stops when the stow mode is selected. The spring in the DCV assembly drives the valve actuator piston and moves the DCV to the stow direction. The directional valve override linkage lets the valve turn without the stop rod moving. Air is admitted to the air motor. The ballscrew turns and the ballnut and ballscrew actuator begin moving toward stow. When the actuator is about .25 inch from fully stowed, the stop rod moves the DCV toward neutral. When closed, the DCV has bleed air holes which allows air to drive the CDU to the full stow stop to pre--load the actuation system.

Deploy Operation Air from the DPV moves the valve actuator piston to the DEPLOY position. The helix rod turns the DCV as the valve actuator piston moves. The piston and pivoted lever release the cone brake, and the air motor rotates turning the ballscrew in the deploy direction. The ballnut and ballscrew actuator move to deploy. The stop rod is pulled toward the deploy stop as the actuator approaches fully deployed. At about 1.5 inches from full deploy, the stop rod touches the ballnut. The stop rod then moves the DCV to the neutral position to stop airflow to the air motor, and engage the cone brake. The stop rod also activates the switches in the CDU position indicating switch module. This causes the T/R PRSOV to close and controls indication of thrust reverser position.

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ENGINE EXHAUST

Figure 10 SCL

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May -- 2001

CDU Operation Page: 21

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

ROTARY FLEXIBLE DRIVE SHAFT General The flexible drive shafts transmit power from the center drive unit to the upper and lower angle gearboxes. There are two drive shafts on each reverser half, one for each angle gearbox. The CDU turns both flexible drive shafts. Access is through the fan cowl. Each drive shaft has an outer casing with mounting flanges and a drive shaft core. The outer casing is corrosion resistant steel lined with teflon. The drive shaft core is stranded wire. The end of the drive shaft at the CDU is a 3/8 inch spline. The angle gearbox end is a 0.2 inch square shaft. The two drive shafts are different lengths. Maintenance Practices if removal is required, open the fan cowl panel. Release the CDU brake. Open the quick release clamps securing the shafts to the thrust reverser torque box. NOTE:

LOWER RH AND UPPER LH DRIVE SHAFTS HAVE TWO CLAMPS, LOWER LH AND UPPER RH DRIVE SHAFTS HAVE ONE CLAMP.

PRECAUTIONS SHOULD BE TAKEN TO PREVENT CORE FROM SLIDING OUT OF CASING. ANY CONTACT WITH UNCLEAN SURFACES WILL REQUIRE CORE REPLACEMENT Remove bolts and washers securing the shaft to the CDU and the angle gearbox. Remove the complete flexible drive shaft unit. CAUTION:

IF ONE FLEXIBLE DRIVE SHAFT ON REVERSER HALF FAILS, BOTH SHAFTS ON THAT HALF MUST BE REPLACED AS TORSIONAL LIMITS MAY HAVE BEEN EXCEEDED. Installing the flexible drive shaft is the opposite of removal. A thrust reverser actuation system rigging procedure must be done after installing a flexible drive shaft.

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CAUTION:

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ENGINE EXHAUST

Figure 11 SCL

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May -- 2001

Rotary Flexible Drive Shaft Page: 23

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

ANGLE GEARBOX AND BALLSCREW ACTUATOR General Three ballscrew actuators move the translating cowl. One of the ballscrew actuators is driven directly by the CDU. The other two ballscrew actuators are driven by the angle gearboxes. The gearboxes are driven by the CDU through the flexible drive shafts. Access is through the fan cowl. Each gearbox has two square input drives to connect a rotary flexible drive shaft and to permit manual operation, and a splined output for the ballscrew actuator connection. The square drive opposite the drive shaft end is capped. This end may also be used to lock the actuator or for rigging. The 0.2 inch drive requires a special tool to fit the hole. The gearbox decreases the flexible drive shaft speed by a 3:1 ratio. The ballscrew actuator is coupled to the gearbox spline. A stop collar (not shown) is pinned to the end of the ballscrew to limit actuation length. The ballnut and actuator tube translates as the ballscrew turns. Removal The angle gearbox and ballscrew actuator must be removed as a unit. The angle gearbox can be separated from the ballscrew actuator after removal. To remove, deploy the translating cowl 6--8 inches to access the ballscrew actuator clevis pin. Remove the flexible drive shaft, then the clevis pin, and finally the gearbox and actuator.

For Training Purposes Only

CAUTION:

NOTE:

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ENSURE THAT THE DRIVE SHAFT CORE DOES NOT SLIDE OUT OF OUTER CASE WHEN REMOVING THE ROTARY FLEXIBLE DRIVE SHAFT. DO NOT REMOVE THE CLEVIS PIN RETAINING CLIP BOLT. BACK THE BOLT OUT ONLY ENOUGH TO ROTATE THE RETAINING CLIP. THE NUT PLATE WILL BE DAMAGED IF THE BOLT IS REMOVED.

WHEN INSTALLING A GEARBOX AND ACTUATOR THE SIDE PLATE ON THE GEARBOX MUST BE FACING INWARD.

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Page: 24

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ENGINE EXHAUST

Figure 12 SCL

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May -- 2001

Angle Gearbox and Ballscrew Actuator Page: 25

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

ELECTROMECHANICAL (TRAS) BRAKE General The electro--mechanical brakes (also called the thrust reverser actuation system or “TRAS” brake) provide a third level of safety to prevent uncommanded deployment of the thrust reversers in flight. (The auto stow system, the locking center drive units, and the TRAS brakes provide three levels of safety.) The brake mechanism has a separate, dedicated electrical circuit for its control that is independent of other thrust reverser components.

For Training Purposes Only

Description There are two electro--mechanical brakes installed on each engine, one on each thrust reverser half. The brakes are mounted on brackets attached to the fan reverser torque boxes. Each brake is connected to its upper angle gearbox by a flexible drive shaft. The electro--mechanical brakes are solenoid activated disk brakes. When 28 V dc is applied to the brake solenoids, the brakes will release to permit thrust reverser operation. These brakes lock their reverser half by locking the flex drive cable at the upper actuator. Operation The electro--mechanical brake (TRAS lock) is spring loaded to the fully braked position. Dual rotors contacting stators provide the braking force friction. To release the brake, the solenoid is energized by electrical current from the thrust reverser actuation system relays and switches. This solenoid force acts against the springs to reduce the rotor / stator friction force, thus releasing the brake. A manual lockout lever is mounted to the upper surface of the brake. Lifting of this lever will cause an internal cam to act against the springs to reduce the rotor / stator friction force, thus releasing the brake. The lockout lever is used during manual extension of the translating cowl for maintenance and rigging of the thrust reverser. The lockout manual release handle will automatically be returned to the brake position when the fan cowl is closed.

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May -- 2001

Page: 26

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ENGINE EXHAUST

Figure 13 SCL

JGB

May -- 2001

Electro-- Mechanical Brake - TRAS Page: 27

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Figure 14 SCL

JGB

May -- 2001

TRAS Locked and Solenoid De-- energized Page: 28

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Figure 15 SCL

JGB

May -- 2001

TRAS Unlocked Solenoid Energized Page: 29

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ENGINE EXHAUST

Figure 16 SCL

JGB

May -- 2001

TRAS Manual Released Page: 30

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

NOTES :

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May -- 2001

Page: 31

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

CDU -- POSITION FEEDBACK TRANSDUCER General One CDU position feedback transducer is mounted to the upper auxiliary drive pad on each CDU. Each transducer unit has two electrical connectors. One goes to EEC channel A and one to EEC channel B. The feedback transducer unit has a bearing mounted driveshaft, a reduction gearbox, and two rotary variable differential transformers (RVDT‘s). The drive shaft is turned by the CDU while the thrust reverser deploys and stows. The output of the drive shaft is reduced through the gearbox and is applied to a single rotor shaft common to both RVDT‘s. The rotor shaft rotates through a 77 ° arc when the translating cowl is deployed, and returns to its original position when the translating cowl is stowed. There is a viewing window on the opposite end of the feedback transducer unit from the drive shaft. The window is for rigging the sensor in the stow position. The RVDT‘s convert the angular position of the rotor shaft into electrical signals that are read by the EEC. Each RVDT receives an excitation from the EEC and returns two position signals to the EEC. The EEC reads the return signals in terms of percent--of--deployment. A reading of 100 % indicates full deployment (rotor shaft displaced 77° ). A reading of 0 % indicates the translating cowl is fully stowed and that the rotor shaft is at the rig point. The operational range of the input to the EEC is from --5 to 105 %. The EEC uses the translating cowl position information to control the thrust reverser interlock actuator. Indications The EEC sends the thrust reverser position information to EICAS. The information appears on the EPCS maintenance page next to the thrust reverser left (T/R L) and thrust reverser right (T/R R) headings. If the EEC is not able to sense thrust reverser position due to a failure in the CDU position feedback transducer, the EICAS status and maintenance message L (R) ENG REV POS appears. This message is latched.

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May -- 2001

Page: 32

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ENGINE EXHAUST

Figure 17 SCL

JGB

May -- 2001

CDU Position Feedback Transducer Page: 33

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

CDU POSITION INDICATING SWITCHES The CDU position indicating switch assembly gives thrust reverser position indication and controls the T/R PRSOV solenoid to stow or deploy the thrust reverser. One switch assembly is installed on each CDU. An electrical cable goes from the switch assembly, along the torque box, to a bracket near the top of the reverser half. Access is through the fan cowl panels. Each switch assembly has a deploy switch and a stow switch. The switches are double--pole double--throw type switches. The switch assembly is a line replaceable unit. During removal, the spring and washer can fall out. The spring is tapered, with the large end going into the CDU housing when it is installed. A line replaceable electrical cable connects the CDU position indicating switch assembly to the airplane wiring harness. The cable on the right thrust reverser half (not shown) has two electrical connectors. One goes to the CDU position indicating switch assembly, the other goes to the airplane wiring harness. The cable on the left thrust reverser half has four electrical connections. One goes to the airplane wiring harness, one to the CDU position indicating switch assembly, and one each to the DPV and its pressure switch.

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RVDT POSITION INDICATION

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May -- 2001

Page: 34

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ENGINE EXHAUST

Figure 18 SCL

JGB

May -- 2001

CDU Position Indicating Switches Page: 35

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST LEVER INTERLOCK ACTUATOR General The thrust reverser interlock actuators (one per engine) prevent the movement of reverse thrust power levers above idle until the translating cowls are at least 60 % deployed. The interlock actuators also prevent the movement of the forward thrust power levers above idle until the translating cowls are at least 80 % stowed. The interlock actuators are located below the autothrottle assembly in the forward equipment center.

For Training Purposes Only

Operation The interlock actuator has a reversible motor that operates on 28 V dc power. The motor extends or retracts a linear actuator. The linear actuator is connected to the autothrottle quadrant. The EEC controls the extension and retraction of the actuator. When the actuator is retracted and the reverse thrust lever is stowed, the forward thrust lever can be advanced. When the actuator is extended, the reverse thrust levers can be advanced from the reverse idle detent.

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May -- 2001

Page: 36

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Figure 19 SCL

JGB

May -- 2001

Thrust Reverser Interlock Actuator Page: 37

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER INTERLOCK ACTUATOR Operation When the reverse thrust lever is raised to the reverse idle detent position, the translating cowls start to deploy. When the CDU position feedback transducers indicate that both translating cowls are deployed 60 % or more, the EEC provides a ground to the T/R interlock relay. 28 V dc power goes to the T/R interlock actuator motor extend windings, and the actuator extends. Power is removed from the motor when the actuator is fully extended. When the actuator is extended, the reverse thrust can be increased. When the reverse thrust lever is lowered to the stowed position, the translating cowls start to stow. When the CDU position feedback transducers indicate that both translating cowls are deployed 20 % or less, the EEC removes the ground to the T/R interlock relay. 28 V dc power then goes to the T/R interlock actuator motor retract windings, and the actuator retracts. Power is removed from the motor when the actuator is fully retracted.

For Training Purposes Only

Indications If the interlock actuator does not move to the fully retracted position when the thrust lever angle is greater than 43° (about 10° of forward thrust lever movement) for more than 10 seconds, an EICAS status and maintenance message L (R) REV INTERLOCK appears. The message is latched.

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May -- 2001

Page: 38

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ENGINE EXHAUST

Figure 20 SCL

JGB

May -- 2001

Thrust Reverser Interlock Actuator Operation Page: 39

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER CONTROL SWITCHES Three thrust reverser control switches control the electrical signals to deploy or stow the thrust reverser. The control switches are in the pilots control stand (P8). One switch, in the forward thrust lever handle, controls the signal to the T/R PRSOV. The other two switches, in the microswitch pack assembly, control the signals to the electro--mechanical brakes (TRAS brakes) and to the DPV. The T/R PRSOV switch closes when the reverse thrust lever is raised more than 10° . The DPV control switch closes when the reverser thrust lever is raised above 29° . This signals the directional pilot valve to open, directing air to the DEPLOY side of the CDU air motor. At 29° the TRAS lock switch closes, providing power to several relays which unlock the electro--mechanical brakes and signal the T/R PRSOV to open.

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ENGINE EXHAUST

SCL

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May -- 2001

Page: 40

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

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ENGINE EXHAUST

Figure 21 SCL

JGB

May -- 2001

Thrust Reverser Control Switches Page: 41

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER GENERAL OPERATION General Thrust reversers are used by the flight crew to decelerate the airplane immediately after landing. Normal thrust reverser operation requires that the airplane be on the ground, engine running, fire switch in normal, and both pneumatic pressure and electrical power available. Deploy Operation The thrust reverser deploys as the reverse thrust lever is raised to a detent position called the reverse idle detent. (The reverse idle detent is a hard mechanical stop controlled by the interlock actuator.) The flight crew pulls the reverse thrust levers to the reverse idle detent position just after the airplane is safely on the runway after landing. This movement is usually a quick continuous motion to the stop. Movement to the reverse idle detent stop covers about 35° of angular motion. There are four solenoids in the thrust reverser activation system (TRAS). All four solenoids must be energized to deploy the thrust reversers. As the reverse thrust lever is pulled back through 10° of motion a switch in the power lever assembly enables power to the thrust reverser sequencing relay (K2184) through the center drive unit (CDU) position switch module. As the reverse thrust lever is pulled back through 29° of motion two switches in the autothrottle microswitch pack assembly are closed. One of these switches energizes the directional pilot valve (DPV) solenoid. The DPV opens, but no air is yet available. The other switch powers relays that energize the two solenoid operated electro--mechanical brakes (TRAS brakes) . One brake is located on each half of the thrust reverser. Additional relays in the circuit to the electro--mechanical brake solenoids energize the T/R sequencing relay which provides power to the T/R PRSOV solenoid. Air is now available to the DPV and to the CDU.

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May -- 2001

The DPV provides sense (control) air to both CDU for that engine. With the DPV open, the two CDU unlock. The CDU air motors drive their internal ballscrew actuators and their upper and lower angle gearbox and ballscrew actuators through flexible drive shafts. The translating sleeves are deployed by the three ballscrew actuators. The translating sleeves pull the blocker doors down to shut off the normal path of fan air. Fan air is forced to exit through the deflectors. The deflectors provide a change in direction of the fan air which results in a deceleration of the airplane. The T/R PRSOV, is de--energized when the translating sleeves are fully deployed. The DPV and the TRAS brake solenoids remain energized.

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ENGINE EXHAUST

Figure 22 SCL

JGB

May -- 2001

Thrust Reverser General Components Page: 43

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

Engine Operation During the approach to landing, the engine is not permitted decelerate below flight idle. After touchdown, the engine speed is maintained at flight (high) idle for 5 seconds by a time delay relay on the engine discretes card. This allows 5 seconds for the pilot to decide to go around or to use reverse thrust. If the pilot does neither, after 5 seconds the engine will decelerate to ground (low) idle and the crew will use the airplane brakes to slow down. If the pilot selects reverse thrust, the engine will not be permitted to operate above reverse high idle speed until both halves of the thrust reverser have deployed. A physical stop is provided by the interlock actuator which prevents the reverse thrust levers from being moved past this detent position. When the electronic engine control receives a feedback signal that the thrust reverser translating sleeves are both near their full deployment, the interlock actuator is permitted to release the detent stop for the reverse thrust levers. At that time the crew can pull the levers back to the full reverse position. The engine can then accelerate to full reverse thrust power. Thrust reverser deployment and engine acceleration to full reverse thrust usually takes less than 5 seconds total time. When the airplane has slowed down to about 60 knots, the flight crew will move the reverse thrust levers forward to the stow position. The DPV will close causing the center drive unit motors to operate the ball screw actuators to stow the thrust reversers. Stow Operation When the crew pushes the reverse thrust levers forward to the stow position, the 29° switches and the 10° switch open. The T/R PRSOV opens to provide air to the CDU’s. The DPV closes. The electro mechanical locks are free. The CDU air motor stows the translating sleeves and blocker doors. When the thrust reversers are fully stowed, all solenoids de--energize. The T/R PRSOV closes and the electro--mechanical brakes lock the upper ball screw actuators.

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May -- 2001

Thrust Reverser Indications When both halves of a thrust reverser are fully deployed, a green REV indication will appear on the upper EICAS display just above the N1 digital display. When both of the translating sleeves are fully stowed there is no REV message shown. When either or both of the translating sleeves are between the fully stowed and fully deployed position, a yellow REV indication appears above the N1 indication. No thrust reverser messages are shown to the flight crew in flight unless there is an actual abnormal inflight deployment of a thrust reverser. Then the yellow or green REV indication could be observed. After the airplane has been on the ground for 60 seconds, faults in the thrust reverser system detected in--flight will illuminate the REV ISLN light and cause the EICAS advisory and latched maintenance message “L (R) REV ISLN VAL” to be displayed. Appearance of these indications on the ground (the messages and the light are inhibited in--flight by air / ground logic) mean either: -- that the reverser may not deploy when commanded on the ground, or -- that the thrust reverser relay module (TRRM) detected and latched an in--flight fault in the reverser system. Thrust Reverser Relay Module The thrust reverser relay module (M1987) (located in the main equipment center) monitors operation of the thrust reverser system. If in--flight faults lasting more than 5 seconds occur, magnetically latched relays will illuminate light emitting diode indication lights on the module’s front panel The thrust reverser relay module provides fault indications for both engines. It incorporates a self test and a lamp test capability.

Page: 44

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ENGINE EXHAUST

Figure 23 SCL

JGB

May -- 2001

Thrust Reverser General Operation Page: 45

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER OPERATION - ELECTRICAL Operational Description -- Electrical Circuits The electrical control system consists of four switches, four solenoids, two position switches, and eight relays for each thrust reverser. Operation of the left engine thrust reverse will be explained. The operation of the right engine thrust reverser is the same, but the components have different numbers and locations. Deploy Mode For an engine thrust reverser deployment the T/R PRSOV, DPV and the two TRAS solenoids all must be energized. To energize the four solenoids, the airplane must be on the ground. With the forward thrust levers at the forward idle position the pilot rotates the reverse thrust lever aft. Rotation of the reverse thrust lever to the rear sequentially closes three switches: -- the T/R control switch (S5), -- the T/R DPV control switch (S11), -- and the TRAS lock switch (S21). The T/R control switch (S5) is the first to close at approximately 10° of reverse thrust lever rotation. At approximately 29° of reverse thrust lever rotation the T/R DPV control and the TRAS lock switches close. The DPV solenoid, T/R sequence relay (K2184), and TRAS lock release relay (K2182) are energized; followed by the T/R PRSOV solenoid (V360), the left and right TRAS solenoids, and the T/R unstow relay (K26); and finally the TRAS lock release control relay (K2188). The proper sequencing of the four controlling solenoids is critical. The DPV solenoid is the first to be energized even though it is controlled by one of the 29° switches. The T/R PRSOV solenoid and the left and right TRAS solenoid are essentially energized simultaneously, however, the TRAS brakes are released prior to pneumatics being available to drive the CDU‘s . There is approximately a 160 millisecond window between the TRAS brake release and the CDU‘s spinning up to speed thereby insuring that the TRAS brakes are not released under load. With proper sequencing, the engine thrust reverser, driven by the CDU‘s, translates to the fully deployed position.

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May -- 2001

Stow Mode During stow operations, the reverse thrust levers are moved forward and down. There is no stop position between deployed and stowed. The 29° switches open first and then the 10° switch opens. The DPV closes. The T/R PRSOV opens to drive the translating sleeves to the stow position. Position switches signal the T/R PRSOV to close, removing air from the CDU‘s. Two seconds after removal of the pneumatic operating pressure from the thrust reverser system, the 28 V dc power is removed from the electro--mechanical brake solenoids and the brakes engage again.

Page: 46

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ENGINE EXHAUST

Figure 24 SCL

JGB

May -- 2001

Thrust Reverser Electrical Operation Page: 47

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER INDICATING SYSTEM OPERATION General This system gives indications of thrust reverser position and malfunctions. No thrust reverser messages are shown to the flight crew in flight unless there is an actual abnormal in--flight deployment of a thrust reverser. Then the yellow or green REV indication could be observed. T/R Position Indication When both halves of a thrust reverser are fully deployed, a green REV indication will appear on the upper EICAS display just above the N1 digital display. When both of the translating sleeves are fully stowed there is no REV indication shown. When either or both of the translating sleeves are between the fully stowed and fully deployed position, a yellow REV indication appears above the N1 indication.

The thrust reverser relay module will latch a fault in any of the following conditions exist for more than 5 seconds while the airplane is in--flight: -- An unstowed sleeve is detected by the limit switches on the center drive unit. The LED labeled RESTOW COMMAND will be illuminated. -- The electro--mechanical brake solenoids are being commanded to release the brakes due to power being present at the thrust reverser activation system (TRAS) lock release control relay (K2188). The LED labeled TRAS UNLOCK will be illuminated. -- Pneumatic pressure is present downstream of the T/R PRSOV as indicated by the pressure switch mounted on the directional pilot valve. The LED labeled PRSOV PRESSURE will be illuminated.

For Training Purposes Only

T/R Malfunction Indications After the airplane has been on the ground for 60 second, faults in the thrust reverser system detected in--flight will illuminate REV ISLN light and cause the EICAS advisory and latched maintenance message ”L (R) REV ISLN VAL” to be displayed. Appearance of these indications on the around (the messages and the light are inhibited in--flight by air / ground logic) mean either: -- that the reverser may not deploy when commanded on the ground, or -- that the thrust reverser relay module (TRRM) detected and latched an in--flight fault in the reverser system. Thrust Reverser Relay Module The thrust reverser relay module (M1987) (located in the main equipment center) monitors operation of the thrust reverser system. If in--flight faults lasting more than 5 seconds should occur, magnetically latched relays will illuminate light emitting diode indication lights on the module’s front panel. The thrust reverser relay module provides fault indications for both engines. It incorporates a self test and a lamp test capability. The thrust reverser relay module only monitors the reverser system while the airplane is in the air mode. It is inhibited on the ground. However, the TRRM can be utilized to monitor the reverser system on the ground to aide troubleshooting by pushing the test enables switch located on the front panel. A reset switch releases the magnetically latched relays to turn off the fault lights. A lamp test switch illuminates all light emitting diodes while pressed. SCL

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May -- 2001

Page: 48

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ENGINE EXHAUST

Figure 25 SCL

JGB

May -- 2001

Thrust Reverser Indicating System Operation Page: 49

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

TRANSLATING COWL MANUAL DEPLOY / STOW This procedure covers manually deploying and stowing the translating cowl using either a manual speed wrench or an air--powered wrench. Do not extend a translating cowl with the thrust reverser open. WARNING:

FAILURE TO FOLLOW THIS INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT. REFER TO 27--61--00/201 FOR APPROPRLATE SPOILER / SPEEDBRAKE DEACTIVATION PROCEDURE. INADVERTENT SPOILER MOVEMENT CAUSED BY ACTUATING THRUST LEVERS COULD RESULT IN SERIOUS INJURY TO PERSONNEL.

CAUTION:

DO NO DEPLOY THRUST REVERSER IF THRUST REVERSER COWL IS OPEN. DAMAGE TO THE TRANSLATING COWLS AND STRUT WILL OCCUR. BE SURE AREA AFT OF THRUST REVERSER IS CLEAR OF ALL EQUIPMENT, WORKSTANDS, ETC. DAMAGE WILL RESULT IF THRUST REVERSER COLLIDES WITH EQUIPMENT.

CAUTION:

WHEN TRANSLATING THRUST REVERSER MANUALLY, WATCH FOR SINGLE ACTUATOR AND CDU OPERATION. IF THIS OPERATION SHOULD OCCUR, STOP TRANSLATING THRUST REVERSER AND CHECK FOR UNINSTALLED OR BROKEN FLEXIBLE DRIVE SHAFTS.

CAUTION:

IF AIR--POWERED WRENCH IS USED TO DEPLOY OR STOW THRUST REVERSER TRANSLATING COWL, WATCH FOR FEEDBACK ROD MOTION WHEN NEARING FULL DEPLOY/STOW. WHEN MOTION IS DETECTED, REMOVE AIR--POWERED WRENCH AND COMPLETE CYCLE WITH A MANUAL SPEED WRENCH. CENTER DRIVE UNIT WILL LOCK UP WITH EXCESSIVE TORQUE.

CAUTION:

ENSURE LOCKOUT PLATE SQUARE DRIVE IS VISIBLE WHEN PLATE IS INSTALLED ON CDU MANUAL DRIVE PAD. THRUST REVERSER WILL FAIL TO OPERATE IF PLATE IS IMPROPERLY INSTALLED.

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May -- 2001

Deploy Open the circuit breakers to remove power from the T/R PRSOV. Turn off the spoiler/speedbrake control system, and put a DO--NOT--OPERATE identifier on the reverser thrust lever. Open the fan cowl panels and release the CDU manual brake. Unlock the thrust reverser electromechanical brake. A manual lockout lever is mounted to the upper surface of the brake. Lifting of this lever will act against the springs to reduce the rotor/stator force, thus releasing the brake. The lever has been designed to be pushed back to the braked position on closing the fan cowl. Remove the lockout plate to expose the manual drive. Turn the drive pad with a 1/4 inch square drive air wrench or manual speed wrench to deploy the translating cowl. Invert and reinstall the lockout plate to deactivate the CDU. Stow Unlock the CDU manual brake (if locked) and remove the lockout plate. Insert the 1/4 inch square drive wrench. Press the CDU stow rig indicator plunger and turn the drive pad to stow the translating cowl. Stop turning when the rig indicator plunger moves further inward and begins to move back out. If necessary, reverse the direction of the wrench (toward deploy) to find the bottom of the rig indicator plunger motion. Check that translating cowl is fully stowed by observing the position of the rig indicator plunger through the CDU rig window. The rig indicator plunger must be seated in the groove of the extension tube flange. Return the CDU manual brake handle to the locked position and install the lockout plate so that the square extension is visible. Make sure that the translating cowl is properly rigged by checking that the gap between the torque box and translating cowl is between .060 and .150 inch.

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Figure 26 SCL

JGB

May -- 2001

Translating Cowl Manual Deploy / Stow Page: 51

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BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

TRANSLATING COWL POWER DEPLOY / STOW USING EXTERNAL AIR DIRECTLY TO THE CDU General This procedure covers power translation of the translating cowl using a ground pneumatic air source connected directly to the CDU. Do not extend a translating cowl with the thrust reverser open beyond the 34° (first stick) position. WARNING:

For Training Purposes Only

WARNING:

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FAILURE TO FOLLOW THIS PROCEDURE COULD RESULT IN INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT. REFER TO 27--61--00/201 FOR APPROPRIATE SPOILER / SPEEDBRAKE DEACTIVATION PROCEDURE. INADVERTENT SPOILER MOVEMENT CAUSED BY ACTUATING THRUST LEVERS COULD RESULT IN SERIOUS INJURY TO PERSONNEL. ENSURE REVERSE THRUST LEVERS ARE IN THE FORWARD (STOWED) POSITION AND THRUST REVERSER CONTROL CIRCUIT BREAKERS ARE OPENED. INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT COULD OCCUR WHEN PROVIDING EXTERNAL PNEUMATIC POWER. THRUST REVERSER WILL DEPLOY WHEN THE REVERSE THRUST LEVERS ARE MOVED AFT TO REVERSE IDLE POSITION. ENSURE AREA AFT OF THRUST REVERSER IS CLEAR OF PERSONNEL AND EQUIPMENT BEFORE OPERATING THE THRUST REVERSER. INJURY TO PERSONNEL AND / OR DAMAGE TO AIRPLANE MAY OCCUR. WITH PNEUMATIC POWER PROVIDED, DEPLOYED THRUST REVERSER STOWS IF ELECTRICAL POWER IS LOST TO DIRECTIONAL PILOT VALVE. FAILURE TO DEACTIVATE THRUST REVERSER FOR GROUND MAINTENANCE COULD RESULT INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT.

May -- 2001

WARNING:

THRUST REVERSER WILL STOW WHEN REVERSE THRUST LEVERS ARE MOVED FORWARD TO FORWARD IDLE POSITION. ENSURE PERSONNEL AND EQUIPMENT ARE CLEAR OF THRUST REVERSER BEFORE REVERSER OPERATION. INJURY TO PERSONNEL AND / OR DAMAGE TO AIRPLANE MAY OCCUR. DO NOT DEPLOY THRUST REVERSER TRANSLATING COWLS WHEN THE THRUST REVERSER OPEN BEYOND THE 34° POSITION. DAMAGE TO THE TRANSLATING COWLS AND STRUT WILL OCCUR. ENSURE AREA AFT OF THRUST REVERSER IS CLEAR OF ALL EQUIPMENT, WORKSTANDS, ETC. DAMAGE RESULTS IF THRUST REVERSER COLLIDES WITH EQUIPMENT. ENSURE EXTERNAL PNEUMATIC POWER SOURCE SUPPLIES CLEAN AND DRY AIR TO CENTER DRIVE UNIT. FOREIGN OBJECTS AND MOISTURE COULD IMPAIR OPERATION.

Deploy First, open selected circuit breakers on the P5 panel and install DO--NOT-CLOSE identifiers; see MM 78--31--00. Next deactivate the spoiler / speedbrake control system, ensure the reverse thrust levers are in the forward (stow) position, ensure that the thrust reverser is not open beyond the 34° position, ensure that the core cowl panels are removed or closed. Open the fan cowl. Remove the blue cap opposite the CDU pneumatic supply and connect pneumatic power from a ground air source. Slowly pressurize to 20--30 psig. Remove the DO--NOT--CLOSE identifiers and close the T/R PRSOV circuit breakers; see MM 78--31--00. Place the reverse thrust levers to the reverse idle position and allow translating cowl to fully deploy. Stow Provide pneumatic power and place reverse thrust lever to forward (stow) position. Allow translating sleeve to fully stow. Reduce pneumatic pressure to zero and disconnect ground pneumatic source. Install, tighten and lockwire the blue cap on the CDU air connection. Ensure the thrust reverser is fully stowed by checking that the gap between the torque box and the translating cowl is 0.060 -- 0.150 inch at the center drive unit. Return the aircraft to normal.

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Figure 27 SCL

JGB

May -- 2001

Translating Cowl Operation with External Air Source Page: 53

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ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

TRANSLATING COWL POWER DEPLOY / STOW USING PRSOV HOLD--OPEN EQUIPMENT General This procedure covers power translation of the thrust reverser translating sleeve using air from the opposite engine, external pneumatics applied to the airplane, or APU pneumatic power. This air in the pneumatic ducting can not normally flow back through the PRSOV to the T¡R PROV. This procedure uses manual opening of the PRSOV and locking it open with hold open equipment. The procedure for using the hold open equipment is in AMM 36--11--09/201.

For Training Purposes Only

WARNING:

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JGB

FAILURE TO FOLLOW THIS PROCEDURE COULD RESULT IN INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT. REFER TO 27--61--00/201 FOR APPROPRlATE SPOILER/ SPEEDBRAKE DEACTIVATION PROCEDURE. INADVERTENT SPOILER MOVEMENT CAUSED BY ACTUATING THRUST LEVERS COULD RESULT IN SERIOUS INJURY TO PERSONNEL. ENSURE REVERSE THRUST LEVERS ARE IN THE FORWARD (STOWED) POSITION AND THRUST REVERSER CONTROL CIRCUIT BREAKERS ARE OPENED. INJURY TO PERSONNEL AND/OR DAMAGE TO EQUIPMENT COULD OCCUR WHEN PROVIDING EXTERNAL PNEUMATIC POWER. THRUST REVERSER WILL DEPLOY WHEN THE REVERSE THRUST LEVERS ARE MOVED AFT TO REVERSE IDLE POSITION. ENSURE AREA AFT OF THRUST REVERSER IS CLEAR OF PERSONNEL AND EQUIPMENT BEFORE OPERATING THE THRUST REVERSER. INJURY TO PERSONNEL AND / OR DAMAGE TO AIRPLANE MAY OCCUR. WITH PNEUMATIC POWER PROVIDED, DEPLOYED THRUST REVERSER STOWS IF ELECTRICAL POWER IS LOST TO DIRECTIONAL PILOT VALVE. FAILURE TO DEACTIVATE THRUST REVERSER FOR GROUND MAINTENANCE COULD RESULT IN INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT. May -- 2001

WARNING:

THRUST REVERSER WILL STOW WHEN REVERSE THRUST LEVERS ARE MOVED FORWARD TO FORWARD IDLE POSITION. ENSURE PERSONNEL AND EQUIPMENT ARE CLEAR OF THRUST REVERSER BEFORE REVERSER OPERATION. INJURY TO PERSONNEL AND / OR DAMAGE TO AIRPLANE MAY OCCUR.

CAUTION:

DO NOT DEPLOY THRUST REVERSER TRANSLATING COWLS WHEN THE THRUST REVERSER IS OPEN BEYOND THE 20 ° POSITION. DAMAGE TO THE TRANSLATING COWLS AND STRUT WILL OCCUR. ENSURE AREA AFT OF THRUST REVERSER IS CLEAR OF ALL EQUIPMENT, WORKSTANDS, ETC. DAMAGE RESULTS IF THRUST REVERSER COLLIDES WITH EQUIPMENT. ENSURE EXTERNAL PNEUMATIC POWER SOURCE SUPPLIES CLEAN AND DRY AIR TO CENTER DRIVE UNIT. FOREIGN OBJECTS AND MOISTURE COULD IMPAIR OPERATION.

Deploy First, open selected circuit breakers on the P11 panel and install DO--NOT-CLOSE identifiers; see AMM 78--31--00. Next deactivate the spoiler/speedbrake control system, ensure the reverse thrust levers are the forward (stow) position, ensure that the thrust reverser is not open beyond the 34° position, ensure that the core cowl panels are removed or closed. Open the fan cowl. Install the hold--open equipment on the PRSOV. Provide pneumatic power to the airplane. Remove the DO--NOT--CLOSE identifiers and close the T/R PRSOV circuit breakers; see AMM 78--31--00. Place the reverse thrust levers to the reverse idle position and allow translating cowl to fully deploy. Stow Provide pneumatic power and place reverse thrust lever to forward (stow) position. Allow translating sleeve to fully stow. Ensure the thrust reverser is fully stowed by checking that the gap between the torque box and the translating cowl is 0.060 -- 0.150 inch at the center drive unit. Return the aircraft to normal.

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Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 28 SCL

JGB

May -- 2001

Translation Cowl Operation By PRSOV Hold Open Device Page: 55

Lufthansa LAN Technical Training

ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

TRANSLATING COWL POWER DEPLOY / STOW USING THE GROUND SERVICE SWITCH General This procedure covers power translation of the thrust reverser translating sleeve using air from the opposite engine, external pneumatics applied to the airplane, or APU pneumatic power. This air in the pnematic ducting can not normally flow back through the PRSOV to the T/R PRSOV. This procedure electrically opens the PRSOV using the ground service switch. The ground service switch is a guarded, normally open, spring--loaded off switch that is located next to the top of the oil tank.

For Training Purposes Only

WARNING:

SCL

JGB

FAILURE TO FOLLOW THIS PROCEDURE COULD RESULT IN INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT. REFER TO 27--61--00/201 FOR APPROPRIATE SPOILER / SPEEDBRAKE DEACTIVATION PROCEDURE. INADVERTENT SPOILER MOVEMENT CAUSED BY ACTUATING THRUST LEVERS COULD RESULT IN SERIOUS INJURY TO PERSONNEL. ENSURE REVERSE THRUST LEVERS ARE IN THE FORWARD (STOWED) POSITION AND THRUST REVERSER CONTROL CIRCUIT BREAKERS ARE OPENED. INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT COULD OCCUR WHEN PROVIDING EXTERNAL PNEUMATIC POWER. THRUST REVERSER WILL DEPLOY WHEN THE REVERSE THRUST LEVERS ARE MOVED AFT TO REVERSE IDLE POSITION. ENSURE AREA AFT OF THRUST REVERSER IS CLEAR OF PERSONNEL AND EQUIPMENT BEFORE OPERATING THE THRUST REVERSER. INJURY TO PERSONNEL AND / OR DAMAGE TO AIRPLANE MAY OCCUR. FAILURE TO DEACTIVATE THRUST REVERSER FOR GROUND MAINTENANCE COULD RESULT IN INADVERTENT THRUST REVERSER OPERATION WITH POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT.

May -- 2001

WARNING:

THRUST REVERSER WILL STOW WHEN REVERSE THRUST LEVERS ARE MOVED FORWARD TO FORWARD IDLE POSITION. ENSURE PERSONNEL AND EQUIPMENT ARE CLEAR OF THRUST REVERSER BEFORE REVERSER OPERATION. INJURY TO PERSONNEL AND / OR DAMAGE TO AIRPLANE MAY OCCUR.

CAUTION:

DO NOT DEPLOY THRUST REVERSER TRANSLATING COWLS WHEN THE THRUST REVERSER IS OPEN BEYOND THE 34° POSITION. DAMAGE TO THE TRANSLATING COWLS AND STRUT WILL OCCUR. .

Deploy First, open selected circuit breakers on the P11 panel and install DO--NOT-CLOSE identifiers; see MM 78--31--00. Next deactivate the spoiler/speedbrake control system, ensure the reverse thrust levers are in the forward (stow) position, ensure that the thrust reverser is not open beyond the 34° position, ensure that the core cowl panels are removed or closed. Open the fan cowl. Provide pneumatic power to the airplane; see MM 36--00. Push the applicable L or R ENG 0FF switch--lights on the air supply module on the P5 panel to the open position. Remove the DO--NOT--CLOSE identifiers and close the T/R PRSOV circuit breakers; see MM 78--31. Place the reverse thrust levers to the reverse idle position. Lift the guard on the PRSOV ground service switch. Push the switch up to the on position and hold it there. Allow the translating cowls to fully deploy. Release the ground service switch. Stow Provide pneumatic power. Push the applicable L or R ENG 0FF switch light on the air supply module on the P5 panel to the open position and place reverse thrust lever to forward (stow) position. Lift the guard on the PRSOV ground service switch. Push the switch up to the ON position and hold it there. Allow the translating sleeves to fully stow. Release the groundservice switch. Ensure the thrust reverser is fully stowed by checking that the gap between the torque box and the translating cowl is 0.060 -- 0.150 inch at the center drive unit. Return the aircraft to normal.

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Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 29 SCL

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May -- 2001

Translating Cowl Operation By Ground Service Switch Page: 57

Lufthansa LAN Technical Training

ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER DEACTIVATION AND LOCKOUT General This procedure covers steps to deactivate the thrust reverser for ground maintenance and mechanically lock the reverser for flight dispatch. Deactivation WARNING:

WITH PNEUMATIC POWER PROVIDED, DEPLOYED THRUST REVERSER WILL STOW IF ELECTRICAL POWER IS LOST TO DIRECTIONAL PILOT VALVE CAUSING POSSIBLE INJURY TO PERSONNEL AND / OR DAMAGE TO EQUIPMENT.

THIS PROCEDURE IS FOR GROUND INADVERTENT THRUST REVERSER TRANSLATION MAY OCCUR IF PROCEDURE IS USED TO DEACTIVATE THRUST REVERSER FOR FLIGHT DISPATCH. First, open the circuit breakers on the P12 panel to remove power from the T/R PRSOV. Put DO--NOT--OPERATE identifiers on the reverse thrust levers. Open the fan cowl panels. Remove, invert and reinstall the lockout plates on both CDU‘s and attach REVERSER DEACTIVATED pennants. CAUTION:

For Training Purposes Only

Lockout To lockout the thrust reverser for flight dispatch, deactivate the CDU‘s as for ground maintenance. Make sure the translating cowls are fully stowed so that the holes in the three translating cowl brackets line up with the holes in the torque box flange. The cowl can be manually stowed by removing the unused drive pad cover on the angle gearbox and turning the gearbox. Remove the six locking bolts and the three red DO NOT OPERATE plates that are stored on the torque box. Reinstall the plates on to the flange, locking the translating cowl in place with the bolts.

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Page: 58

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 30 SCL

JGB

May -- 2001

Thrust Reverser Deactivation and Lockout Page: 59

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER COWLS OPENING SYSTEM To open and close the thrust reverser halves, there are two hydraulic cowl opening actuators (one for each half). The thrust reverser power pack supplies pressurized hydraulic fluid to the actuators through hydraulic lines. The pack uses dc power from the ground handling bus. TO control the opening system, there is a T/R door control switch on the left and right side of the engine fan case. The T/R check valve is a safety device used during closing. There is one hold open rod for each thrust reverser half. A backup for the thrust reverser power pack is the handpump.

For Training Purposes Only

Lufthansa LAN Technical Training

ENGINE EXHAUST

SCL

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Page: 60

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 31 SCL

JGB

May -- 2001

Thrust Reverser Opening System Page: 61

Lufthansa LAN Technical Training

ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER POWER PACK AND CONTROL SWITCHES Thrust Reverser Power Pack The thrust reverser power pack supplies pressurized hydraulic fluid to the thrust reverser opening actuators. The power pack contains: -- A hydraulic reservoir. -- An electric motor to supply power to a hydraulic pump. -- Two valves to control the flow of hydraulic fluid to the opening actuators. -- Two solenoids that control the valves. The power pack uses 28 V dc to operate the motor and to control the valves. If a leak occurs in the power pack, a drain line drains the fluid into the oil tank scupper drain line. Maintenance Practices If the fluid level is low, remove the fill port dipstick and add fluid up to the dipstick full mark.

For Training Purposes Only

T/R Door Control Switches The two T/R door control switches control the electric motor and the two valves in the thrust reverser power pack. The switches are on the fan case at the 5:00 and 7:00 position. Each switch has three positions: -- up, -- stop -- down.

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Page: 62

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 32 SCL

JGB

May -- 2001

Thrust Reverser Power Pack and Control Switches Page: 63

Lufthansa LAN Technical Training For Training Purposes Only

ENGINE EXHAUST

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

THRUST REVERSER COWLS OPENING / CLOSING SYSTEM OPERATION General There are two methods of opening and closing the thrust reverser cowls: -- electrical operation -- hand pump operation. Before opening the thrust reverser: -- Make sure the leading edge slats are retracted and deactivated. -- Deactivate the thrust reverser for ground maintenance. -- Open the fan cowls.

Hand Pump Operation If electrical power is not available, or in case of motor / pump failure, a hand pump can be used to open the cowls. The hand pump connections are the quick disconnects located on the fan case at the 5:00 and 7:00 positions.

Electrical Operation Open the three reverser tension latches and the latch ring assembly for the cowl (s) being opened. To open the left cowl (the right cowl procedure is the same), move and hold the left T/R door control switch to the “UP” position. This energizes the left power pack relay to supply 28 V dc to the motor. Moving the switch to ’up’ also energizes solenoid ’A’ which opens the left power pack valve. Hydraulic fluid flows to the left opening actuator. A pressure relief valve in the power pack permits hydraulic fluid to flow back into the reservoir if the cowl is jammed or latched. The electrical circuit prevents opening both cowls at the same time. Hold the switch ’up’ until the thrust reverser is fully open. Release the switch. The switch is spring--loaded to ’stop’. This turns off the hydraulic pump. Solenoid ’A’ remains energized. Extend and install the thrust reverser hold open rod. Push the switch to ’down’. This de--energizes solenoid ’A’. Push the T/R check valve plunger in. This permits hydraulic fluid to flow back to the power pack reservoir. The thrust reverser hold open rod now holds the weight of the cowl. To close the left cowl, move the left T/R door control switch to ’up’. This removes the load on the hold open rod. Remove and stow the hold open rod. Move the switch to down. Push the T/R check valve plunger in. This permits hydraulic fluid to flow back to the reservoir. A restrictor in the line keeps the cowl from closing too guickly.

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Page: 64

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 33 SCL

JGB

May -- 2001

Thrust Reverser Cowls Operation Page: 65

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

ANGLE GEARBOX

For Training Purposes Only

Lufthansa LAN Technical Training

ENGINE EXHAUST

CENTER DRIVE UNIT

Figure 34 SCL

JGB

May -- 2001

Thrust Reverser Rig Setting Page: 66

Lufthansa LAN Technical Training

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

For Training Purposes Only

ENGINE EXHAUST

Figure 35 SCL

JGB

May -- 2001

Thrust Reverser Indicating Summary Page: 67

BOEING -- 767 / 300 CF6 -- 80C2 78 -- 00

NOTES :

For Training Purposes Only

Lufthansa LAN Technical Training

ENGINE EXHAUST

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May -- 2001

Page: 68

Lufthansa LAN Technical Training For Training Purposes Only

ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

EPCS EEC Monitoring System EEC Inputs / Outputs EPCS PIMU PIMU Panel Description PIMU Operation EEC Flight Fault Recording PIMU Flight Fault Recording PIMU BITE EEC No--Flight Fault Recording PIMU Powering PIMU BITE Procedure PIMU Maintenence Recall EICAS -- EPCS Pages FAULT EXAMPLE Engine Control Logic Messages ARINC 429 Format EPCS ARINC Analysis EPCS Conversion Table Table Two Label 270 Table Two Label 271 Table Two Label 272 Table Two Label 273 Table Two Label 274 Table Two Label 275 Table Two Label 276

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002 004 006 008 010 012 014 016 018 020 022 024 028 030 032 039 041 043 044 045 046 047 048 049 050 051

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Lufthansa LAN Technical Training

ENGINE ENGINE PROPULSION CONTROL SYSTEM EEC MONITORING SYSTEM General The Electronic Propulsion Control System (EPCS) is a full authority, digital electronic, propulsion control system. The EPCS consists of the Electronic Engine Control System (EECS) on each engine, and the airframe elements which interface with them.Each of the two engines incorporates a dual channel, independently--powered, self--checking and automatic--fault--accommodating digital Electronic Engine Control (EEC). The EPCS elements include the thrust lever mechanisms, the thrust reverser position indication, and the interfaces with other systems, such as the Engine lndicating and Crew Alerting System (EICAS), the two Air Data Computers (ADC), the Thrust Management Computer (TMC) and the Standby Engine Indicator (SEI).Two tiers of EPCS maintenance condition indications are provided. The first tier consists of general status and maintenance messages that are shown on the display units. The second tier consists of fault messages displayed on the Propulsion Interface and Monitor Unit ( PIMU ). These messages assist to maintenance personnel in isolating system faults to a specific Line Replaceable Unit (LRU) or to the interfacing circuits between LRUs.The EICAS message PIMU indicates that the PIMU self--test has detected a fault.The PIMU records and stores faults from the EEC. The PIMU receives EEC fault data for a five second period via separate channel A and channel B ARINC 429 data buses either automatically when the air/ground relay closes upon landing, or manually by actuation of the GND TEST switch on the face of the PIMU with an engine running or with ground power applied to the EEC. With ground power applied to the EEC, some systems that require engine rotation will be inoperative. Fault data received is stored in non--volatile memory of the PIMU. Stored fault data may be manually retrieved for visual display on the face of the PIMU.

For Training Purposes Only

FAULT DETECTION LOGIC * Definition The purpose of the FADEC/EEC fault reporting system is to identify the faulty LRU‘s and the type of failure in the control system.

BOEING 767 / 300 CF6 -- 80C2 77 -- 35 * Circuit Checks Output: -- Command circuit checks are only made in the controlling channel. These checks are made while the engine operates or by using ground test power. -- A WRAPAROUND message indicates a failure in the controlling channel. The failure can be in a torque motor, solenoid or wiring harness. Input : -- Circuit checks are made in both channels. These checks are made while the engine operates or by using ground test power. -- A RANGE message indicates that an input signal is below or above the permitted limits or that it does not change at the permitted rate. -- A CROSS--CHECK message indicates that a channel‘s parametric or position input differs from the other channel‘s input by more than the permitted amount. WHEN THE GROUND TEST POWER IS USED THE N1 AND N2 SPEEDS ARE NOT TESTED. -- A FEEDBACK FAILED message indicates a range or cross--check failure in a components feedback circuit. The failure can be in a LVDT, resolver, RVDT or wiring harness.

NOTE:

* Position Checks -- Position Checks are only while the engine operates. The tests are not made during the engine start to prevent incorrect indications. . An EXTERNAL WRAPAROUND message indicates that a system‘s valve is not in its commanded position. The position of the valve is sensed by a valve operated switch. . ATRACK--CHECK message indicates that the system‘s actuator or valve is not in its commanded position. NOTE:

A TRACK--CHECK FAULT CAN BE INDICATED ONLY IF WRAPAROUND, FEEDBACK AND RANGE CHECK TESTS HAVE NOT FOUND ANY FAULTS.

* No Message -- If there is a parameter shift with no message, the cause of that shift was not found by the control system. SCL

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Page: 2

Lufthansa LAN Technical Training

ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

Servo TLA

Command Analog Signal

L EICAS SEI R EICAS

Feedback

TMC

L PIMU

EEC

Engine Sensors

ARINC 429 FMC

For Training Purposes Only

Feedback

DFDAU

L AIR DATA R AIR DATA

Figure 36 SCL

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Jul -- 2002

FADEC Data Flow Page: 3

Lufthansa LAN Technical Training

ENGINE ENGINE PROPULSION CONTROL SYSTEM EEC OUTPUT SIGNALS * Torque Motor Commands -- Calculated by the Central Processor Unit (CPU). -- In a 8 bit format converted by gate array. -- Pulse Width Modulated (PWM) signal converted by torque motor drivers. . Plus or minus current to torque motors on their respective actuators. * Torque Motor Wrap--Around (T/M W/A) -- Torque motor current measured at the output of the torque motor drivers. . Current output converted to a proportional DC voltage. . DC voltage is a “wrapped--around” to the input of the torque motor wrap--around multiplexer. * Torque Motor DC Wrap--Around Selection and Conversion. -- T/M W/A Mux switches the DC voltage input to its output upon command from the CPU. -- Analog to digital multiplexer at the aproximate time, ( as determinated by the CPU ), switches in the selected DC voltage into a 16 bit digital value for use by the CPU. * Torque Motor Wrap--Around Tests -- Measured torque motor current (converted digital value) is range checked against stored upper and lower limits. -- Measured torque motor current is compared to the command value. -- Measured torque motor current is compared to its previous value and the rate of change is checked aginst limits for exceedance.

For Training Purposes Only

EEC INPUT SIGNALS

BOEING 767 / 300 CF6 -- 80C2 77 -- 35 * Rate Tests -- Input signals once converted to digital form within the FADEC/EEC are compared with the previous sample of this input. Changes which exceed known sensor or sensed quantity rate of change capabilities indicate faulty data. Rate tests are effective in detecting intermittent or noisy signals. * Cross--Check -- Test performs a comparision between the local and cross--talked inputs. If either input is failed for any other reason, the test is not performed. Test failure indicates that one or both sensing devices are transmitting incorrect signals. * Fault Latching -- Accommodation of faulty inputs sometimes required control mode reconfiguration to alternate modes of use of similar but not identical input signals. Intermittent fault, if not properly handled can cause cycling between modes or signals, resulting in unacceptable operation. -- The FADEC/EEC contains software techniques called fault latching, which detect intermittent fault conditions and prevent mode cycling. The logic is designed to allows short--term faults to occur, then recover, but to latch out persistent faulty data and annunciate these persisttent faults for maintenance action. The characteristics of this logic are: . Sustained failures lasting longer than the specified fault latch time will cause fault annunciation and exclusion of this input signal. . Short--term failures will cause mode changes, but complete recovery will occur when the fault “healts” except transition to reversionary mode. . Latches are cleared when control is depowered or reset via the fuel cutoff reset.

* Range Tests -- Input signals once converted to digital form within the FADEC/EEC are compared with range limits which represent the valid output extremes of a sensor or the extremes of the sensed quantity. Electrical inputs to the FADEC/EEC are, in general, biased such as that opens or shorts in external wiring will drive the digital signal out of range.

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

INPUTS

OUTPUTS

For Training Purposes Only

COMMAND FROM CHANNELS A OR B

Figure 37 SCL

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Inputs - Outputs Page: 5

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

ELECTRONIC PROPULSION CONTROL SYSTEM EPCS During engine operation, the engine control system constantly tests itself. When a fault is detected, the appropriate discrete bit (or bits) is set in the serial digital data bus discrete data word. For ground test purposes, you can supply power for the EEC when you do not operate the engine. You will find two power switches, one for each engine, on EEC maintenance module found on the P61 panel. You supply the power to the EEC for the ground test when you move the EEC MAlNT L (R) POWER switch from NORM to TEST. With the switch in the test position, you can see the EEC 1 or 2 GND PWR status message on the EICAS STATUS page. In this mode many bits are not set by the EEC, since certain functions do not operate without engine rotation.Each EEC channel on each engine will detect anomalies and hard faults in its processor, its inputs and its outputs. In some cases the exact faulty unit in the system cannot be totally isolated by automatic means; only the particular ”LRU loop” can be flagged. The ”LRU loop” ( command and feeedback signal) includes the EEC interface circuitry, wiring from the EEC to the sensor, servo or other components, and the interfacing component itself. Status or maintenance messages that show on the display units are directly related to fault messages displayed on the PIMU. For example, the EICAS message L (R) ENG CONTROL corresponds to PIMU messages as shown in FIM 73--21--00/101. In most cases, when the above EICAS message is displayed its corresponding PIMU message will be present when the PIMU is interrogated. Fault codes identifv a particular EPCS component malfunction or a related circuit problem. Included in the system are possible internal EEC failures. Each fault is assigned as unique PIMU message.

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Lufthansa LAN Technical Training

ENGINE ENGINE PROPULSION CONTROL SYSTEM

SCL

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Page: 6

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

E1-- 2 OR E1--3

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

Figure 38 SCL

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Jul -- 2002

PIMU Location Page: 7

Lufthansa LAN Technical Training

ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

PROPULSION INTERFACE MONITOR UNIT SYSTEM

DESCRIPTION

General The propulsion interface monitor units (PIMU‘s) provide storage and display of maintenance information from the EEC. Each PIMU also receives and retransmits engine data to the thrust management computer (TMC) and to the digital flight data acquisition unit (DFDAU). There are two PIMU‘s, one for each engine, located in the main equipment center. The left engine PIMU is in the E1--3 rack and the right engine PIMU is in the E2--4 rack.The 115 V ac ground service bus powers the PIMU. The PIMU has two major functions:

PIMU Displays and Controls The PIMU front panel has 3 rows of 8 character light emitting diode (LED) displays, giving it 24 alphanumeric characters for fault display. Each of the 24 characters is formed by a matrix of 35 point LEDs (5 dots wide by 7 dots high).Front panel control switches and an input from the air / ground system control the recording and display capabilities of the PIMU. Automatic recording of flight faults transmitted from the EEC to the PIMU occurs for both EEC channels during the 5 seconds following a landing signal from the air / ground system relays.

Interface Functions The PIMU receives 71 ARINC 429 data words from each channel of the EEC. This data is then transmitted by the PIMU on three output data buses. The TMC receives channel A and channel B data on separate data buses. The DFDAU receives data only for the EEC channel in control on the third data bus. 9 of the 71 data words sent by the EEC contain fault information. The other 62 data words contain engine parameter data. The PIMU also electrically isolates the EEC from the TMC and the DFDAU.

Indications When EEC faults are stored in the PIMU NVM, a discrete signal to EICAS causes the maintenance message “L (R) PIMU” to be displayed. The EEC sends the same 71 words of data to the EICAS computers on another data bus. Some of the faults detected by the EEC and encoded in the 9 ARINC 429 fault words will cause EICAS alert, status, and / or maintenance messages. When these messages appear, the fault isolation manual will call for the -- PIMU BITE-- PIMU GROUND TEST-- PIMU MAINTENANCE RECALL and procedures to be performed. The PIMU description and operations, maintenance practices, and installation procedures are in AMM 77 -- 35 -- 00. The fault isolation manual procedures for using data stored in the PIMU are in chapter 71--PIMU MESSAGE INDEX.

For Training Purposes Only

Monitor Functions Each EEC continuously monitors any faults detected by its BITE. These faults are coded into the 9 ARINC 429 fault words transmitted by the EEC to the PIMU. The EEC will store these faults into its non volatile memory (NVM) when the engine is shutdown. The PIMU can also store these faults in its own NVM, but only when commanded to do so. The PIMU also has BITE capabilities to test its own internal operations.

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Page: 8

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BOEING 767 / 300 CF6 -- 80C2 77 -- 35

For Training Purposes Only

ENGINE ENGINE PROPULSION CONTROL SYSTEM

Figure 39 SCL

JGB

Jul -- 2002

PIMU System Page: 9

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

PIMU Panel Description There is a separate monitor unit for each engine. The monitor unit’s front panel has the following features: a) A 24 character alphanumeric display in three eight character lines, to annunciate stored faults. The top line is used to annunciate the associated channel A or channel B fauIt. Example : 352--14/CH A. The second and third line are used to annunciate an alphanumeric fault label. b) A BIT switch that initiates the memory recall function of the unit and annunciation of the first of any discrete faults stored in memory. Subsequent depression and release of the switch initiates annunciation of additional discrete fault data, one at a time. c) A MONITOR VERIFY switch when depressed causes all segments of the alphanumeric display to illuminate. Release of the switch initiates a seíf--test.The seíf--test verifies the operating integrity of the monitor unit without actually receiving data from active ARINC 429 inputs. d) A GND TEST switch verifies operation of the selected CH A or CH B data bus and the monitor unit’s associated ARINC 429 receiver by receiving and decoding ARINC 429 inputs.e) A RESET switch when depressed wiII cause all stored faults to be cleared and the display to go dark immediately. The RESET switch is guarded to prevent inadvertent operation. f) A MAlNT RECALL switch, when pushed in, will cause the EEC to enter in the maintenance recall mode. In this mode, each subsequent push in and release of the MAl NT RECALL switch wilI cause the EEC to transmit one fault message stored in the memory of the EEC and also the flight Ieg of that fault. The flight leg is calculated by the EEC. To toggle between the display of the recalled fault message and its flight Ieg, push in the BIT SWITCH. To exit this mode, select the MONITOR VERIFY switch.

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Jul -- 2002

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

ALPHANUMERIC DISPLAY

BIT SWITCH MAINT RECALL SWITCH

MONITOR VERIFY

For Training Purposes Only

RESET SWITCH

GROUND TEST CHANNEL SELECTOR

Figure 40 SCL

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PIMU Panel Description Page: 11

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

OPERATION The PIMU is operational when 115 volts ac power is supplied to the PIMU. The system begins functioning when the air/ground relay switches to the ground position. The unit performs a self--test that verifies the intergrity of the monitor unit and the operation of both data buses. Upon successful completion of the selftest the unit receives fault data for a five second period and stores faults in a non--volatile memory. lf the monitor fails the self--test, an indication is sent to the display unit and the monitor turns off. During the five second operating period the monitor unit receives low speed ARINC 429 serial digital data from the channel A and channel B of the corresponding EEC. The data received by the monitor unit contains EEC SYSTEM FAULT bits.A fault is identified by a single data bit. A specific fault associated with a bit is defined by the location of the bit in the word. The words received by the ARINC 429 receiver must contain 32 bits and pass a parity test before the words are read and recorded. The fault data must be received two times consecutively during the sample period before being stored as a fault. Once a fault has been stored in non--volatile memory it can be cleared by operation of the RESET switch. Before recalling the faults stored in memory, the monitor unit can be checked by pressing the MONITOR VERIFY switch. With the switch depressed, all segments of the alphanumeric display are illuminated. Release of the switch initiates a self--test which verifies the operating integrity of the monitor unit without actually receiving data from the ARINC 429 inputs.lf the self--test takes longer than three seconds to accompíish the message TEST IN PROGESS wiIl be displayed until the test is complete. Successful completion of the self-test is annunciated by displaying READY message for ten seconds. Failure of the test is annunciated by displaying PIMU MONITOR FAIL for ten seconds.The ARINC 429 receiver and data bus may also be tested by actuation of the GND TEST switch. The ground test wiIl verify the operation of the selected channel A/B data bus and ARINC 429 receiver.The verification is accomplished by successfulíy receiving and decoding the data inputs. Data is received by actuating the GND TEST switch with an engine running or with ground power applied to the EEC. Ground test failure wilI cause the message CH A or CH B DATA BUS INOP to be displayed for ten seconds.

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Jul -- 2002

After verifying the operation of the data bus, the unit is enabled for a five second period to store any fauIt data received during the test of the selected channel. The applicable CH A or CH B TEST IN PROGRESS message is displayed while the unit is enabled. After the five second period the unit wiII turn off.To recall faults stored in memory the BIT switch is depressed and released. On release the monitor unit turns on and annunciates the first of any fauIts stored in memory. The annunciation of fauIts are in the form of alphanumeric messages. The message corresponding to each fault is displayed on the second and third line of the alphanumeric display. The first line will display the actual label and channel designator.For example: 354 13--A or 155 22--B.The message is displayed on the screen until the BIT switch is depressed again. All channel A faults messages are displayed followed by all channel B faults messages.After the Iast fauIt message has been displayed, pressing the BIT switch wiII cause the monitor to display the message END for ten seconds. At the end of this ten seconds period the display wiIl automatically blank.Depression of the RESET switch will cause both the RAM and the non--volatile memory of the PIMU to be cleared. The PIMU can operate as a device to recall the fault messages and their related flight Ieg number stored by the EEC into the EEC non--volatile memory. This is accomplished when you operate the Maintenance Recall Switch as described above. During the usual operation, the fault messages stored by the EEC should be the same as the messages recorded by the PIMU at each landing. However, it may be useful to recall the EEC non--volatile memory contents if the PIMU memory was cleared by the accidental push in of the PIMU RESET switch. It also may be useful to recall EEC non--volatile memory contents if after an engine start the scheduled flight is aborted before takeoff. In that case, the PIMU wilI not contain fault intormation associated with the aborted flight because the air/ground relay stayed in the ground state.

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353-- 14/CH B

Figure 41 SCL

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EEC FAULT MONITORING AND RECORDING DURING FLIGHT OPERATIONS General To understand how the PIMU BITE operates, it is necessary to understand the way that the EEC and the PIMU manage their respective memory. The PIMU is able to store up to 144 faults per channel for the past 128 flight legs, but the EEC can store no more than 40 faults per channel for the past 64 flights. The 40 fault storage positions, 6 are dedicated to either NO DISPATCH faults or SHORT TERM DISPATCH faults, the remaining 34 storage positions are dedicated to LONG TERM DISPATCH faults or faults that DO NOT AFFECT DISPATCH. EICAS displays the L (R) ENG CONTROL advisory message if a NO DIS PATCH fault is stored in the EEC. The status messages L (R) ENG EEC C1 or L (R) ENG EEC C2 are displayed for SHORT TERM and LONG TERM DISPATCH faults respectively. The L (R) PIMU maintenance message is the only message displayed for faults that DO NOT AFFECT DISPATCH. The EEC NVM can not be erased using the PIMU or any other onboard method. If there are more faults than the memory can hold, the oldest fault will be overwritten by the newest fault detected.

A flight leg is determined by the EEC without inputs from the air / ground systems on the airplane or inputs from any other airplane system. The EEC uses its own sensors to compute a Mach number and altitude. A new flight leg is determined by the EEC when its computed Mach number is roughly equal to 100 knots, or the PO sensor shows a decreased pressure equal to an altitude increase of 400 feet, or the pressure has decreased to an equivalent altitude of more than 16,500 feet (which is higher than any airport on this planet). From that time any faults detected since the EEC‘s latest power--up will be recorded against a new flight leg 1. The detected faults will be transmitted on the data buses to the PIMU, but will be held in an EEC buffer until N2 goes below 20% during engine shutdown. The previous flight leg 1 will then become flight leg 2 in the EEC memory.

Definition of a Flight Leg The EEC monitors all faults from the time that the EEC is powered. It will not store any of these faults into permanent memory until N2 has been above 30% and then goes under 20%. If it is necessary to monitor faults during engine motoring, which does not usually go above 30% N2, the PIMU GROUND TEST switch must be placed to the CH A position momentarily, wait 10 seconds, and then placed in the CH B position while the engine is still motoring. Faults detected by the EEC will be sent to the PIMU on the data buses, but will not be recorded in the EEC memory. The PIMU GROUND TEST procedure may be used to place the faults into PIMU NVM.

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Jul -- 2002

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Figure 42 SCL

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EEC Faults during Flight Operations Page: 15

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

PIMU AUTOMATIC FAULT RECORDING DURING FLIGHT OPERATIONS General PIMU automatic fault recording occurs when the air / ground relay system signals that the airplane has landed. For a period of 5 seconds, the PIMU records in non volatile memory (NVM) any faults being sent over the channel A and the channel B data buses from the EEC. The flight is not finished at the time of landing. Thrust reverser, taxi, and engine shutdown operations are yet to happen. The EEC will continue to monitor the systems for faults. Any faults will be held in the EEC fault buffer until the N2 speed decreases below 20% on engine shutdown. Faults detected by the EEC after touchdown will not be stored by the PIMU. The only way to determine if faults were stored in the EEC NVM after landing is to perform the PIMU maintenance recall procedure. Unless there was an EICAS message that was not appropriate for the results of a normal PIMU BITE procedure, there would not be any indication that hidden faults exist in EEC memory.

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

Figure 43 SCL

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

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PIMU BITE -- MOST RECENT FLIGHT General For 5 seconds after landing, the PIMU automatically records any EEC faults for the current flight in non volatile memory (NVM). The flight leg is not completed at the time of recording of the faults. The EEC NVM will have a record of any faults detected during the reverse thrust, taxi and shutdown phases of the flight. These faults can only be recalled by using the maintenance recall procedure. If faults are stored in the PIMU, an EICAS maintenance message L (R) PIMU appears. Operation Make sure that the 115 V ac ground service bus is powered. First, push the MONITOR VERIFY switch and hold it in. A matrix of point light emitting diodes (LED‘s), 5 LED‘s wide by 7 LED‘s high should appear for each of the 24 character positions. Note if any are not operating, but continue the test. Next, release the MONITOR VERIFY switch. The PIMU enters a self test mode. If the test takes more than 3 seconds, the message TEST IN PROGRESS appears. The message READY appears for 10 seconds if the test was successful. Next, push the BIT switch. The first channel A fault (if any) will appear. To see the next fault, push the BIT switch again. After all of the channel A faults have been shown, the next push of the BIT switch will show the first channel B fault (if any). When all of the faults have been shown, or if there were no faults, the message END appears for 10 seconds. After another 10 seconds the display will blank. Be sure to erase fault data from the PIMU by pushing the RESET switch. This will erase PIMU NVM faults but will not erase the faults that are stored in the EEC. If the PIMU memory is not erased, the faults from this flight will be included with those of the next flight in the PIMU NVM.

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Figure 44 SCL

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PIMU BITE - Most Recent Flight Page: 19

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

EEC FAULT MONITORING AND RECORDING NO FLIGHT CONDITIONS Re--ejected Takeoff In case of a re--ejected takeoff (RTO), the EEC will begin fault monitoring as soon as the EEC is powered. When the engine is started the N2 goes above 30%, so any faults will eventually be stored in non volatile memory. Since the computations made by the EEC will not indicate that the airplane is in the air by : -- Mach number > 100 knots, or -- altitude increase > 400 feet, or -- pressure altitude > 16,500 feet the EEC will not establish a new flight leg 1 during a rejected takeoff. Any faults detected by the EEC will be added to the faults in the existing flight leg 1. The only way to determine what faults may have been stored in the EEC NVM after landing is to perform the maintenance recall procedure. Unless there was an EICAS message that was not appropriate for the results of a normal PIMU BITE procedure, there would not be any indication that hidden faults exist in EEC memory. Since the faults detected during the re--ejected takeoff are included with any faults that might have occurred during the last flight leg 1, it might be difficult to isolate the exact time the faults happened.

Maintenance Ground Runs As with the re--ejected takeoff discussed above, there will not be a new flight leg 1 for maintenance ground runs. There will also not be an air / ground landing signal, so there will not be an automatic PIMU recording. Any faults detected by the EEC will be added to the faults in the existing flight leg 1. The maintenance manual procedures for ground run tests (AMM 7100--00 501 series pages) call for the PIMU ground test procedure before the engine is shutdown. First press the RESET switch to erase the PIMU NVM. Next push the GROUND TEST switch to the CH A position and release it. Wait 10 seconds. Then push the GROUND TEST switch to the CH B position and release it. This will store only the faults detected by the EEC during this ground run in the PIMU NVM. Any faults detected will be stored in the EEC against the latest flight leg 1 along with fault from the other possible ground runs that may have been made since the last takeoff.

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RTO Fault Recording Schedule

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Ground Run Fault Recording Schedule

Figure 45 SCL

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PIMU POWER General The PIMU ground test is used to determine if there are any current faults detected by the EEC. Both the EEC and the PIMU must be powered to conduct two of the three test, because the PIMU has three different mode test. -- PIMU Fault Recall. Only PIMU powered. -- PIMU Ground Test. Both EEC and PIMU powered. -- PIMU Maintenance Fault Recall. Both EEC and PIMU powered. There are three ways to power the EEC: -- put the EEC maintenance switch (P61 panel) to the TEST position -- motor the engine above 11% N2 -- start the engine To supply power to the PIMU, the 115 Vac ground service bus must be powered.

For Training Purposes Only

Operation Test the PIMU by pushing the MONITOR VERIFY switch and releasing it. Wait for the message READY to appear and then go out. A spring loaded return--to off toggle switch on the PIMU starts the test. Push the switch to the CH A position and release. Wait 10 seconds. The message TEST IN PROGRESS appears. The display then blanks. Push the switch to CH B position and release. Wait 10 seconds. The message TEST IN PROGRESS appears. The display then blanks. If a channel is not powered, the message DATA BUS INOP will appear.

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

Figure 46 SCL

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PIMU Power Page: 23

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

PIMU BITE General The PIMU records and stores faults from the EEC. A description of system operation is found in AMM 77--35--00/201. The PIMU message are defined by a label and bit identifier, as example 350--14. The PIMU will show the label and bit that are then correlated to a fault message.The label and bit data can also be used for input monitoring. Examine the wires and connectors to make sure the parts are serviceable. Look for problems in the wires same as breaks or cracks in the wires or the connector covers. Make sure that the ground straps and shields are in good condition and function correctly.

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Figure 47 SCL

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

Figure 49 SCL

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PIMU BITE Sh - 3 Page: 27

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BOEING 767 / 300 CF6 -- 80C2 77 -- 35

PIMU MAINTENANCE RECALL General The maintenance recall procedures allow the recall of the fault history stored in the EEC. Faults from the most recent flight, flight 1, will be displayed first. Then the faults for the next oldest flight that had faults can be shown on the PIMU. This procedure allows us to look at the fault history of that channel of that engine for the last 64 flights. The maintenance recall procedure will transfer faults only for the channel in control. The engine must be shutdown and maintenance ground power applied to the EEC. The faults are brought over from the EEC NVM into the PIMU’s random access memory, one fault at a time. To view the faults that have been recorded in the EEC NVM for the other channel, exit the maintenance recall mode by pushing the MONITOR VERIFY switch, unpower that EEC by cycling the maintenance ground test switch to NORM, then back to TEST, and finally pull the appropriate engine channel circuit breaker. This procedure changes the channel--in--control as shown on the EPCS EICAS page.

To get the faults from the opposite channel, exit the maintenance mode with the MONITOR VERIFY switch, shut off the ground test power, turn the ground test power back on, and pull the appropriate circuit breaker to change the channel in control. The recall procedure for the other channel can then be done.

Operation Push the MONITOR VERIFY switch to test the PIMU. READY will show if there are no faults in the PIMU itself.Pushing the MAINTENANCE RECALL switch begins the transfer of data from the EEC NVM to the PIMU random access memory (RAM), one fault bit at a time. You must wait 5 seconds while TEST IN PROGRESS is shown. When the transfer of the fault is completed, the FLIGHT LEG # message appears. Pushing the BIT switch will display the fault. The dollar ($) symbol between the label and bit designation shows that this is maintenance mode data from the EEC NVM. Only faults for the channel in control will be shown. Pushing the BIT switch again and again will toggle between the fault just seen and the flight leg number. To see the next fault you must push the MAINTENANCE RECALL switch, wait for 5 seconds until the FLIGHT LEG # is shown, and then push the BIT switch to display the fault. The fault isolation manual only requires that the latest flight leg with faults be recalled. For historical data or to analyze recent problems, it may be required to recall all of the faults for all possible 64 flights. A maximum of 40 faults can be recalled for each channel.

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Jul -- 2002

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ENGINE ENGINE PROPULSION CONTROL SYSTEM

Figure 50 SCL

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PIMU Maintenance Recall Page: 29

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

EICAS EPCS EPCS Page 1 AND Page 2 The values for various engine control and status parameters appear on the EPCS maintenance page. The parameters appear as real time, AUTO EVENT or MAN EVENT data. Data from both channels of the EEC on each engine appear. The channel which is currently controlling engine operations (or which was controlling the engine in the case of AUTO EVENT or MAN EVENT) is indicated by a square around the channel letter. Page 2 of the EPCS display is accessed by pressing the EPCS switch a second time. Page 2 is a Real Time page only. There are no Manual Events and no Auto Events. The hexidecimal ARINC 429 labels can be decoded using FIM 71--PIMU MESSAGE INDEX.

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PAGE 2

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PAGE 1

Figure 51 SCL

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EICAS EPCS Pages Page: 31

BOEING 767 / 300 CF6 -- 80C2 77 -- 35

A EXAMPLE : When the fault ocurr, the EPCS detec this fault monitored by the EEC and stored in the PIMU, each fault stored in the PIMU is determined by the corresponding message and a additional L/R PIMU. The next example to following the isolation of the fault, represent a message “L ENG CONTROL” , after the PIMU interrogation you obtain First window 350 $23 Second window TRA SIG FAIL Third DETECTED The FIM take both EICAS message and PIMU interrogation result and address the procedure according with this inputs.

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350 $23A

350 $23B

TRA SIG FAIL

EEC CH--B

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353 $19B SENSE FAIL

Figure 52 SCL

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PIMU BITE Page: 33

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EICAS MESSAGE

OR

PIMU MESSAGE

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NEXT STEP

Figure 53 SCL

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CONFIRM

OR

NEXT STEP

Figure 54 SCL

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FIM Step 2 Page: 35

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FIM Step 3 Page: 36

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FIM Step 4 Page: 37

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FIM Step 5 Page: 38

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Engine Control Logic -- 1 Page: 39

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Engine Control Logic -- 2 Page: 40

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ARINC 429 Aeronautical Radio INCorporated 429 Is the language which permit the computers communication. A knowledge of numbering systems ls fundamental to understanding computers and their operation. All numbering systems are used to count objects or perform mathematical calculations and each is a set of symbols and characters, commonly referred to as digits. The systems normally produce feedback signals which are sended to the computers, but this analogs signals are converted into the computers in digital inputs for to process inside of the CPU, reconverted after the process in analog outputs as command or driver signals.

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Analog Inputs (FDBK)

A D

Central Processing Unit

A D

Analog Outputs (CMD)

Analog signals from system elements are encoded into BCDC data words for transmission. BCD words transmit several numeric characters and discrete signals to using systems. Examples of data transmitted into this word format includes engine information to DFDAU or TMC. The structure of the BCD word format is divided functionally and consists of: -- Label code -- Source/Destination Identifier (SDI) -- Data Field -- Sign Status Matrix (SSM) -- Parity Bit

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HEXADECIMAL BINARY

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BIT NUMBER

Figure 60 SCL

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Standard Convertion Table Binary Hexadecimal Page: 42

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Figure 61 SCL

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EPCS Page 2 Page: 43

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EPCS Convertion Table Page: 44

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Table Two Label 270 Page: 45

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Table Two Label 271 Page: 46

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Table Two Label 272 Page: 47

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Table Two Label 273 Page: 48

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Table Two Label 274 Page: 49

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Table Two Label 275 Page: 50

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Table Two Label 276 Page: 51

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NOTES :

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