B767 ATA 70-80 Student Book
May 3, 2017 | Author: Elijah Paul Merto | Category: N/A
Short Description
B767 ATA 70-80 Tranining Manual. Contains Operation of the GE CF6-80C2 Engine on the B767....
Description
GE CF6-80C2F POWERPLANT CH 71-80
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ATA 71 GE CF6-80 C2F TABLE OF CONTENTS
FUEL FLOW INDICATION................................................................... 68 AIR SYSTEMS GENERAL DESCRIPTION ......................................... 70
TOC CF6-80C2FADEC: ........................................................................ 2 ABBREVIATIONS AND ACRONYMS ................................................... 3 POWER PLANT CF6-80C2F................................................................. 4 ENGINE COWLING............................................................................... 6 THRUST REVERSER ......................................................................... 10 CORE COWL PANELS ....................................................................... 16 ENGINE MODULE CONSTRUCTION................................................. 18 AIRFLOW STATION............................................................................ 20 ENGINE CONFIGURATION................................................................ 22 FAN ROTOR MAINTENANCE ............................................................ 24 ACCESSORY DRIVES MODULE ....................................................... 26 ENGINE COMPONENTS .................................................................... 28 ENGINE BORESCOPE INSPECTION PORTS................................... 32 ENGINE VENTS AND DRAINS........................................................... 34 ENGINE CHANGE............................................................................... 36 ENGINE PRESERVATION.................................................................. 38 OIL DISTRIBUTION SYSTEM OPERATION....................................... 40 LUBE AND SCAVENGE PUMP .......................................................... 44 MAGNETIC CHIP DETECTORS ......................................................... 46 OIL INDICATING SYSTEM ................................................................. 52 OIL INDICATION OPERATION ........................................................... 54 ENGINE FUEL DISTRIBUTION SYSTEM........................................... 56 FUEL PUMP ........................................................................................ 58 FUEL FILTER ...................................................................................... 58 SERVO FUEL HEATER ...................................................................... 60 FUEL NOZZLES .................................................................................. 60
VARIABLE BYPASS VALVES ............................................................. 72 VSV AND VBV CONTROL .................................................................. 76 COMPRESSOR DISCHARGE TEMPERATURE SENSOR (T3)......... 80 CCCV SYSTEM ................................................................................... 84 TURBINE CASE COOLING................................................................. 86 STANDBY ENGINE INDICATOR (SEI) ............................................... 96 ENGINE TACHOMETER SYSTEM ..................................................... 98 ENGINE FUEL AND CONTROL MESSAGES................................... 102 AIRBORNE VIBRATION MONITORING SYSTEM............................ 106 ENGINE N2 SPEED CARDS............................................................. 112 CONDITION MONITORING .............................................................. 114 PROPULSION INTERFACE MONITOR UNIT (PIMU) SYSTEM ...... 116 ELECTRONIC PROPULSION CONTROL SYSTEM (EPCS)............ 126 FADEC SYSTEM DESCRIPTION ..................................................... 128 EEC DISCRETES PRINTED CIRCUIT CARD .................................. 138 HMU FUEL METERING OPERATION .............................................. 143 EEC INPUTS/OUTPUTS ................................................................... 146 CONTROL MODES ........................................................................... 155 ENGINE IDLE SELECT ..................................................................... 158 START SYSTEM AIR SOURCES...................................................... 160 ENGINE IGNITION LEADS, PLUGS AND START CONTROL ......... 166 THRUST REVERSER SYSTEM........................................................ 170 T/R PRESSURE REGULATING AND DIRECTIONAL PILOT VALVE 180 TRANSLATING COWL DEPLOY/STOW........................................... 190 DEACTIVATION AND LOCKOUT .................................................... 196
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ABBREVIATIONS AND ACRONYMS ACC - Active Clearance Control ACTR- Actuator AVM - Airborne Vibration Monitoring CCCV - Core Compartment Cooling Valve CTRL- Control EEC - Electronic Engine Control FADEC - Full Authority Digital Engine Control GE - General Electric gnd - ground hdlg - handling HMU - Hydro-mechanical Unit HP - High Pressure IDG - Integrated Drive Generator LP - Low Pressure PIMU - Propulsion Interface Monitoring Unit PRSOV - Pressure Regulating and Shutoff Valve TAI - Thermal Anti-Ice TIP - Training Information Point T/R - Thrust Reverser T12 - Temperature at Station 1.2 svc - Service VBV - Variable By-pass Valves VSV - Variable Stator Vanes N1 - Low Pressure Compressor Speed N2 - High Pressure Compressor Speed
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GENERAL - POWER PLANT CF6-80C2F Purpose The two strut mounted engines provide the airplane with thrust, electrical power, pneumatic power, and hydraulic power. General Description The power plant system is supported by the airplane strut. This includes the engine, cowling, exhaust, mount and drain components. The General Electric CF6-80C2F engines are a high bypass ratio (see engine specifications), dual rotor, turbofan engine. Engine cowling consists of the inlet cowl, fan cowl and core cowl. The exhaust system discharges fan and turbine air through separate paths to atmosphere. Fan exhaust is directed through a pneumatic thrust reverser. Turbine exhaust passes through the exhaust sleeve. The forward and aft engine mounts carry thrust, vertical, side and torque loads. Specifications CF6-80C2F • • • • • • • •
Rated Thrust Classification 60,000 Pounds Flat Rated Temperature 86F Bypass Ratio 5.15 to 1 Compressor Pressure Ratio 29.9 to 1 EGT Redline (Max) 960C N1 Redline (Max) 117.5% N2 Redline (Max) 112.5% Weight 9485 lbs
TURBINE EXHAUST SLEEVE
CORE COWL PANEL
THRUST REVERSER
1
FAN COWL CHINE (INBOARD SIDE ONLY) FAN COWL PANEL INLET COWL
1
INBD
EXHAUST SYSTEM COMPONENTS SHOWN FOR REFERENCE ONLY
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71-00-C2F-001
1
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GENERAL - ENGINE COWLING Purpose The cowling provides an aerodynamically smooth protective surface over the engine, engine-mounted components, and accessories. The cowling controls airflow around and through the engine, provides access to various areas of the engine case and fan case. General Description The cowling for each engine includes the inlet cowl, fan cowl, thrust reverser and core cowl. Access doors and openings are provided on the cowling to facilitate maintenance and servicing. The turbine exhaust consists of hot, combusted gases exiting the low pressure turbine at high velocity. The major components of the turbine exhaust system are the exhaust sleeve and plug. Fan cowls, thrust reversers and core cowls are mounted to the strut with hinges. Inlet cowl, exhaust sleeve and exhaust plug are bolted directly to the engine case. General Operation The engine cowling opening sequence is fan cowl, thrust reverser, core cowl, and closing sequence is in reverse order. Together with associated exhaust systems, powerplant cowling performs several functions. It minimizes aerodynamic drag of the engine installation. It protects components within from hostile flight environments, provides sound suppression and directs airflow for proper engine operation. Also powerplant cowling provides for fire and over-pressure protection. Inlet Cowl Constructed of aluminum structure, with honeycomb core acoustical lining, and kevlar/graphite external panels. Approximately 106 inches outside diameter, 55 inches long and weighs 564 lbs.
Fan Cowls Constructed of aluminum structure, with nomex honeycomb and kevlar/graphite external panels. The Fan Cowls are approximately 106 inches outside diameter, 53 inches long and weighs a total of 137 lbs. or 68.5 lbs each side. Thrust Reverser Cowls The fan thrust reverser cowls incorporate a self-contained hydraulic system to power open the reverser halves for engine access. They provide the forward thrust duct and also block and redirect this thrust forward to accomplish reverse thrust. The Fan Thrust Reverser Cowls are approximately 104 inches outside diameter, 63 inches long and weighs a total of 1538 lbs. or 769 lbs. each side. Core Cowls The Core Cowl panels are constructed of aluminum, titanium, and cres (corrosion resistant stainless steel). The Core Cowls are approximately 72 inches outside diameter, 59 inches long and weighs a total of 244 lbs. or 122 lbs. each side. Exhaust Sleeve And Plug Both the exhaust sleeve and plug are constructed of welded honeycomb.
STRUT
EXHAUST SLEEVE HINGE (TYP)
HINGE (TYP)
CORE COWL PANELS PRESSURE RELIEF DOOR
OIL TANK ACCESS DOOR
THRUST REVERSER HALVES CHINE FAN COWL PANELS PRESSURE RELIEF DOOR
INLET COWL
NOTE: EXHAUST PLUG NOT SHOWN
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GENERAL - FAN COWL PANELS Purpose The left and right fan cowl panels protect the engine fan case.
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. With the sleeve retracted, engage hold-open rod onto engine mounted bracket and release sleeve. Brackets are mounted on engine flanges.
Access The fan cowl panels are hinged to the strut and fair 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 to provide access to components on the engine fan case. Characteristics Each fan cowl overlaps the corresponding thrust reverser half. A pressure relief door, located midway up the left cowl, opens to relieve excessive fan cowl compartment pressures. The right fan cowl contains an access door to service the engine oil tank without opening the cowl. Two hold-open rods are installed on each fan cowl panel to support the cowl in the open position. The extended hold-open rods engage brackets on the fan case to hold the fan cowl open to positions of 40 or 55 degrees from the bottom centerline. The free ends of the rods are stowed in receivers on the cowl when not in use.
Closing Fan Cowl Panels The corresponding thrust reverser half must be closed before closing the fan cowl panel. Disengage aft hold-open rod first, then disengage forward holdopen rod. Retract sleeve at receiver end of hold-open rod and disengage rod from engine mounted bracket. Rotate and slide collar in direction indicated to unlock hold-open rod from its extended position. UNLOCKED indication should be visible. Retract hold-open rod and engage into fan cowl panel receiver. CAUTION: DO NOT ALLOW FAN COWL PANEL TO SLAM CLOSED. DAMAGE TO FAN COWL PANEL AND/OR ENGINE COMPONENTS MAY RESULT. Push fan cowl panels together and engage latches.
Opening Fan Cowl Panels Release fan cowl latches and engage hold-open rods. Engage forward holdopen rod first, then engage aft hold-open rod.
WARNING: ADEQUATE SUPPORT OF FAN COWL PANEL MUST BE MAINTAINED WHILE ENGAGING HOLD-OPEN RODS TO PREVENT INJURY TO PERSONNEL AND/OR ENGINE COMPONENTS. Retract sleeve at receiver end of hold-open rod to disengage rod from receiver. Fully extend rod to locked position. Check that red UNLOCKED indicator band is not visible.
WARNING: ENSURE THAT HOLD-OPEN ROD IS FULLY EXTENDED AND LOCKED TO PREVENT ACCIDENTAL CLOSING OF
HINGE (3) OIL TANK ACCESS
AFT HOLD - OPEN ROD
SLEEVE RECEIVER
RIGHT FAN COWL PANEL HOLD-OPEN RODS LATCH (3) FORWARD HOLD-OPEN ROD
SLEEVE SECONDARY LOCK
INNER COLLAR
RECEIVER
RED UNLOCKED BAND
INNER SEGMENT OUTER COLLAR
FWD
OUTER SEGMENT SLEEVE
HOLD-OPEN ROD
LEFT FAN COWL PANEL (WITH HOLD-OPEN RODS STOWED)
ENGINE-MOUNTED RECEIVER
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GENERAL - THRUST REVERSER Purpose The thrust reverser, in the stowed position, provides a smooth surface for the fan exhaust to produce thrust. In the deployed position, the thrust reverser redirects the fan exhaust to produce reverse thrust. Access A hydraulic system is used to open each thrust reverser half to access engine components. The thrust reverser halves are attached to the strut and fair with the fan cowl and core cowl. Opening the thrust reverser provides access to components on the high pressure compressor case and accessory gearbox. Characteristics Each thrust reverser half overlaps the corresponding core cowl panel. The thrust reverser half is hinged to the lower part of the strut with three hinges. Thrust reverser halves are latched together with tension latches and the thrust ring latch assembly. The thrust ring latch assembly consists of 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. Operation The inner and outer duct walls provide a flow path for fan air exhaust. Translating cowl, drag links and blocker doors are used to direct fan exhaust through the deflectors when the thrust reverser is deployed. The pneumatically powered center drive unit (CDU) and ball screw actuators move the translating cowl to the deployed position. In the stowed position, the deflectors are covered by the translating cowl reducing drag. The translating cowl is lined with acoustical material for sound suppression.
HINGE (3) UPPER LATCH DEFLECTORS
ANGLE GEARBOX AND BALLSCREW ACTUATOR
INNER DUCT WALL OUTER DUCT WALL
TRANSLATING COWL CENTER DRIVE UNIT
LOWER LATCH THRUST REVERSER TORQUE BOX UPPER AND LOWER LATCH HANDLES
UPPER LATCH CABLE (NOT VISIBLE) STOWED POSITION
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DRAG LINK
ANGLE GEARBOX AND BALLSCREW ACTUATOR DEPLOYED POSITION
THRUST REVERSER B767-3S2F
BLOCKER DOOR
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GENERAL - THRUST REVERSER LATCH ASSEMBLIES Purpose The thrust ring latch assembly secures the outer leading edge of the thrust reverser halves to the aft flange of the fan frame and case. It transmits reverser loads into the engine fan frame instead of the strut hinges.
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.
Location and Access
Adjustment
This assembly is mounted around the leading edge of each thrust reverser half. Access is gained by opening the appropriate fan cowl panel.
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.
Characteristics The upper latch of the mounting ring is a hook that slips into a "U" bolt, mounted to a bracket, on top of the fan stator case. Upper latching force is controlled by the adjustable "U" bolt. The bottom latch is a barrel nut that fits into a "claw" type clevis bracket mounted at the bottom of the fan case. The barrel nut is adjustable to control lower latching force. Upper and lower latch handles are used to open/close upper and lower latches. The upper latch cable is adjustable. The thrust ring latch assembly may be removed if the thrust reverser half is replaced. Operations and Limitations Opening the thrust ring latch assembly requires pulling lower latch handle outward until latch pin bottoms in slot. Rotate upper latch handle outward disengaging latch pin from slot. The upper latch is now disengaged from the "U" bolt. Rotate lower latch handle outward disengaging barrel nut from clevis bracket. Closing the thrust ring latch assembly requires engaging barrel nut with clevis and rotate lower latch handle inward. Rotate upper latch handle inward engaging latch pin in slot. Upper latch should engage "U" bolt. CAUTION: DO A VISUAL CHECK THAT THE LATCH RING HOOK HAS ENGAGED THE "U" BOLT WHEN CLOSING. ALSO, WHEN OPENING THE COWLING ENSURE THE LATCH HOOK IS CLEAR OF THE RING HOOK. FAILURE TO COMPLY WITH THIS COULD CAUSE DAMAGE TO THE COWLING AS WELL AS THE ENGINE PYLON.
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.
FAN STATOR CASE
U-BO LT
UPPER LATCH
THRUST R ING LATCH ASSE MBLY
UPPER LATCH CABLE
FWD LOWER LATCH LATC H BOLT NUT LATCH ANCHOR BOLT FAN STA TOR CASE
OCTAGONAL OFFSET BUSHING
CLEVIS BRACKET
TENSION LATCH
SPRING SCA LE TEST POINT
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BARREL NUT
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GENERAL - THRUST REVERSER OPENING SYSTEM General The thrust reverser cowl opening is done with a hydraulic power opening system. A hand pump is required for opening/closing the thrust reverser. A hand pump can be connected to a quick disconnect to manually open the thrust reverser. Thrust Reverser Opening Actuator The thrust reverser opening actuator is driven by hydraulic pressure to open each thrust reverser half. 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 thrust reverser opening actuator inlet fitting incorporates a restrictor as a safety device limiting the rate of closure. In the event of a hydraulic line rupture or rapid closure, the restrictor provides a minimum 15 second closing cycle. A 25 micron filter at the input fitting protects the restrictor and actuator assembly from fluid contamination. The thrust reverser opening relief valve is for system high pressure relief and is set 4350 - 4500 psig. Thrust Reverser Hold Open Rods Each thrust reverser half has one hold open rod. The rod pivots from a torque box mount under the center drive unit and is held in stowed position with a quick release clamp. The hold open rod consists of an inner rod telescoped inside an outer tube. The hold open rod is held in the telescoped position by a ball lockpin which passes through both inner rod and outer tube through either of two holes. The hold open rod engages a single bracket on the engine fan case and holds the reverser half open to the 34 or 45 degree position depending on which hole is engaged.
AUXILIARY RESERVOIRS
STRUT
FWD OIL TANK
HYDRAULIC CONNECTOR
THRUST REVERSER OPENING ACTUATOR
HAND PUMP
FAN STATOR CASE ROD END
BALL LOCK PIN
HYDRAULIC CONNECTOR
PLUNGER BUTTON
HOLD OPEN ROD
DUST CAP
FWD
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UPPER LATCH HOOK
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GENERAL - CORE COWL PANELS Purpose
Release core cowl latches and engage hold-open rods. Fully extend rod to locked position. Check that red UNLOCKED indicator band is not visible.
WARNING: ENSURE THAT HOLD-OPEN ROD IS FULLY EXTENDED
The left and right core cowl panels protect the turbine case section of the engine. Location & Access
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.
The core cowl panels are hinged to the strut, and fair with the inner barrel of the thrust reverser on the forward edge and rests against the engine exhaust sleeve on the aft edge. Panels are latched together with three flush mounted tension latches at the bottom. The core cowl panels open to allow access to the combustion and turbine cases of the engine.
With sleeve retracted, engage hold-open rod onto engine mounted bracket.
Characteristics
WARNING: ADEQUATE SUPPORT OF CORE COWL PANEL MUST BE
A pressure relief door on the right core cowl panel opens to relieve excessive core cowl compartment pressures. The door is hinged and latched. Two lanyards are used to restrain the door when it is open. Fire shields installed inside the core cowl panels protect them from high temperatures. A hold-open rod installed on each core cowl panel supports the cowl in the open position. The hold-open rod engages a bracket on the engine and is extended to position the cowl open to 50 degrees from the bottom centerline. The free end of the rod is stowed in a receiver on the cowl when not in use. The support rod is telescopic and varialble on some core cowling. Opening Core Cowl Panels The fan cowl panels and thrust reverser must be open before attempting to open the core cowl panels.
WARNING: BE SURE FAN COWL PANELS ARE OPENED AS REQUIRED BY 78-31-00/201 BEFORE OPENING THRUST REVERSER. FAILURE TO FOLLOW 78-31-00/201 COULD RESULT IN INJURY TO PERSONNEL AND/OR DAMAGE TO FAN COWL PANELS, CORE COWL PANELS, AND THRUST REVERSER.
Closing Core Cowl Panels
MAINTAINED WHILE HOLD-OPEN RODS ARE BEING DISENGAGED TO PREVENT INJURY TO PERSONNEL AND/OR ENGINE COMPONENTS. Retract sleeve at receiver end of hold-open rod to disengage rod. Rotate and slide collar in direction indicated and depress secondary lock to unlock hold open rod from its extended position. The hold open rod is now retracted allowing collar to move to its original position. UNLOCKED indication is visible. CAUTION: DO NOT ALLOW CORE COWL PANELS TO SLAM CLOSED. DAMAGE TO PANEL AND/OR ENGINE COMPONENTS MAY RESULT. Stow hold open rod and lower core cowl panel.
FIRE SHIELD (L AND R SIDE)
HINGE (3)
HOLD OPEN ROD LANYARD
PRESSURE RELIEF DOOR (RIGHT SIDE ONLY)
LATCH (3) UNLOCKED INDICATOR
COLLAR
COMPRESSOR REAR FRAME COLLAR RECEIVER BRACKET
SLEEVE
FWD
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SECONDARY LOCK
RIGHT CORE COWL PANEL WITH HOLD-OPEN ROD STOWED
HOLD-OPEN ROD
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SLEEVE
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GENERAL - ENGINE MODULE CONSTRUCTION System Description The CF6-80C2F is a dual spool, axial flow, high bypass ratio turbofan power plant. It has an integrated fan rotor and low pressure compressor (also referred to as a "booster compressor" and a 14 stage high pressure compressor (HPC). The combustor is annular type. A 2-stage high pressure turbine (HPT) drives the high pressure compressor, while a 5-stage low pressure turbine (LPT) drives the fan and low pressure compressor. 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
HIGH PRESSURE TURBINE MODULE
FAN MODULE
CORE MODULE HIGH PRESSURE COMPRESSOR
LOW PRESSURE TURBINE MODULE
FAN ROTOR AND LOW PRESSURE COMPRESSOR
ANNULAR COMBUSTOR
ACCESSORY DRIVES MODULE
ENGINE MODULE CONSTRUCTION B767-3S2F Page - 19
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AERODYNAMIC STATIONS Identification Gas turbine engine manufacturers adhere to Aerospace Recommended Practice (ARP) 755A when assigning aerodynamic station numbers. This standard was developed by the Society Of Automotive Engineers, Inc. and provides performance station identification and nomenclature systems for gas turbine engines. These identifications are referenced by number and alpha characters and relate to both primary and secondary airflow gas paths. The primary airflow path is identified with numbers 0 through 9 and secondary airflow paths are identified with numbers 10 through 19. Any points of measurement between whole numbers is identified in decimal equivalents. The alpha prefix character(s) are used to clarify whether air temperature or air pressure are being measured. They also indicate the manner in which the temperature or pressure is being measured. Of the many characters available those used on the GE engines are: T = Temperature P = Pressure S = Static Engine Instrument Sensor/Station Relationships Temperature and pressure sensors are labeled with a T or a P, and a station number which indicates the location of the sensor in the airflow. The CF6-80C2 sensors (not shown) are: • • • • • • • • •
T12: (Electrical) inlet temperature (2) P14: Fan duct pressure (Condition Monitoring System) P2.5: HPC inlet pressure T2.5: HPC inlet temperature (Condition Monitoring System) P3: Compressor discharge pressure T3: Compressor discharge temperature P4.9: LP turbine inlet pressure (Condition Monitoring System) T4.9: LP turbine inlet temperature (EGT) T5: LP turbine exit temperature (Condition Monitoring System)
FAN DUCT PRESSURE P14
LP TURBINE INLET PRESSURE AND TEMPERATURE P4.9 T4.9
wdmt-h72-00-0001
FAN INLET TEMPERATURE T12
T5 LP TURBINE EXIT TEMPERATURE
P3 T3 COMPRESSOR DISCHARGE PRESSURE AND TEMPERATURE
P2.5 T2.5 HPC INLET PRESSURE AND TEMPERATURE AERODYNAMIC STATIONS B767-3S2F Page - 21
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GENERAL - ENGINE CONFIGURATION General Configuration The basic engine configuration for the CF6-80C engine consists of four Sump location: • • • •
Sump A Sump B Sump C Sump D
Sump A has the #1, 2, and 3 bearings. The B sump has #4, Roller and Ball type bearings. The C sump contains the #5 bearing and is located just forward of the HPT inlet. The D sump is the furthest aft on the engine at the LPT outlet.. The LPC module on the CF6-80C engine has four stages of compression and a single stage fan section. This is also referred to as the booster section. The HPC area consists of 14 stages of compression and is located in the main core of the engine forward of the combustion case. A single annular combustor is used on the engine for fuel introduction and combustion. The HPT consists a two stage turbine and is used to drive the 14 stage HPC. The LPT has a five stage turbine and is used to drive the booster section of the engine.
COMPRESSOR SECTION
COMBUSTION SECTION
TURBINE SECTION
LOW PRESSURE COMPRESSOR L
HONEYCOMB NESTING AREA HIGH PRESSURE TURBINE (2 STAGES) HIGH PRESSURE COMPRESSOR (14 STAGES)
"C" SUMP
LOW PRESSURE TURBINE (5 STAGE)
#5 BEARING "A" SUMP
#3 BEARING
"B" SUMP
"D" SUMP #6 BEARING #2 BEARING #1 BEARING
ROLLER BEARING #4 ROLLERBEARING #4 BALLBEARING
FAN ROTOR BLADE
ENGINE CONFIGURATION B767-3S2F Page - 23
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GENERAL - FAN ROTOR MAINTENANCE Fan Rotor Spinner The fan rotor spinner is mounted to the fan disk by 38 bolts. A sealing ring reduces air leakage around the joint. When installed, the spinner covers the front of the dovetail slots to help hold the fan blades in place. The spinner is balanced separately from the fan rotor before it is mounted. One of the 38 bolt holes is offset to ensure proper alignment of the spinner and the fan disk. Radial captive nuts in the spinner provide mounting locations for fan rotor trim balance screws to make trim balancing the rotor easier. Trim balance weights are used as necessary, but all holes are filled by a balance weight or a screw plug. Fan Rotor Blades The 38 fan rotor blades are mounted in axial dovetail slots in the Fan Disk. The blades are numbered counterclockwise looking aft. Blade position 1 is the second dovetail slot counterclockwise from the spinner bolt hole which is offset. A spring-loaded spacer and keyed retainer prevent forward motion of the blade in the slot. The mid-span shrouds also prevent fore and aft motion of the blades. Removal of the spacer allows the blade to move radially inward. This disengages the mid-span shroud. Balancing weights may be added to the retainer for coarse balancing of the fan rotor. CAUTION: ALL PARTS REMOVED, EXCEPT BOLTS AND NUTS, SHOULD BE MATCHMARKED OR NUMBERED FOR ASSEMBLY IN ORIGINAL ALIGNMENT AND POSITION. USE ONLY APPROVED MARKING MATERIAL. Note:
When removing only one fan blade or opposite blades, it will be necessary to remove the blade retainer, spacer and key from the adjacent blades to allow enough blade movement to disengage the mid-span shroud.
When fan blades are replaced, the minimum allowable clearance between blade tips and the abradable shroud must be maintained.
CAUTION: ALL FIRST STAGE FAN BLADES, RETAINERS/SPACERS MUST BE INSTALLED BEFORE MEASURING BLADE TIP-TO-SHROUD CLEARANCES. Fan Rotor Spinner The fan rotor spinner is made of aluminum 7075 and is black anodized. It is bolted to the fan disk. The spinner is aerodynamically shaped to minimize inlet drag and to deter ice accumulation. Mounting locations are provided for balance weights for precision balancing of the spinner and fan rotor.
SEALING RING 1ST STAGE BLADE (38 LOCATIONS)
FAN ROTOR SPINNER
KEY
DOVETAIL SLOTS
SPACER BOLT MID-SPAN SHROUD FWD
RETAINER WEIGHT
KEY
CLASS
SPACER FAN ROTOR BLADE OFFSET HOLE
STAGE 1 FAN DISK
1 38
2
37 3
BALANCE SCREW SEAL RING CAPTIVE SHANKNUT SPINNER MOUNTING BOLT PATTERN
FWD
FIRE DETECTION - INTRODUCTION FAN / ROTOR MAINTENANCE B767-3S2F Page - 25
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GENERAL - ACCESSORY DRIVE MODULES General Most of the gear driven engine accessories are mounted on, and driven by the accessory gearbox. Refer to the diagram for the pad locations for the following accessories: Forward Side • • • •
Main engine control (Fuel Control Unit) Lube and scavenge pump assembly EEC Control alternator Hydraulic pump
Aft Side • Integrated Drive Generator (IDG) • Pneumatic starter • Fuel pump
PAD 5 LUBE AND SCAVENGE PUMP HORIZONTAL DRIVE SHAFT
PAD 7 HYDROMECHANICAL UNIT
N2 SPEED SENSOR
OIL TUBE BRACKET
PAD 3 HYD PUMP
ACCESS COVER FOR BORESCOPE ROTATION ADAPTER (REF)
PAD 9 PERMANENT MAGNET ALTERNATOR FORWARD SIDE
PAD 8 FUEL PUMP
PAD 4 IDG PAD 6 PNEUMATIC STARTER AFT SIDE NOTE: ACCESSORIES OMITTED FOR CLARITY
ACCESSORY DRIVES MODULE B767-3S2F Page - 27
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GENERAL - ENGINE COMPONENTS Locations The various engine system components are mounted on the engine. The following component locator, broken down by module, is intended as a general orientation to the engine. Component locations given by clock positions are viewed from aft, looking forward. For more details on specific systems, refer to the appropriate chapter. Fan Module Components located in the engine inlet: • Fan rotor: (Immediate access to fan rotor spinner cone, fan rotor blades.) • Electrical T12 sensor: (2:30 and 10:30) Components mounted on the outside of the fan case: • • • •
Oil tank: (3:00) Oil scavenge filter: (4:00) EEC (9:00) Ignition exciters (8:00)
Components mounted in the fan frame (accessible from the aft side of the fan case): • • • • • • •
Forward main engine mount: (12:00) Variable bypass valve system (not shown) 2 VBV actuators: (3:00 and 9:00) 12 variable bypass valves Transfer gearbox: (6:00) Electrical N1 speed sensor: (2:00) Number 1 bearing vibration sensor connector and spare mounting pad (8:00)
Core Module Compressor Stator Case • Accessory gearbox and heat shield
• • • • • • • • •
Variable stator vane system 2 VSV actuators: (3:00 and 9:00) 2 VSV actuation levers (not shown): (3:00 and 9:00) IDG air/oil heat exchanger: (3:30) Main fuel supply hose Fan discharge air manifolds (for core cooling and turbine case cooling) 8th Stage bleed manifold Compressor Rear Frame Fuel tubes (manifold) - 2 igniter plugs (3:00 and 5:00) - HP and LP recoup air manifolds
High Pressure Turbine Module • Active clearance control (ACC) manifold (fan discharge air) • Stage 2 turbine nozzle cooling manifold (11th stage compressor air) Low Pressure Turbine Module Low Pressure Turbine Stator • 8 thermocouple probes • High pressure recoup manifolds (from diffuser) • Active clearance control manifolds (fan discharge air) Turbine Rear Frame • Rear main engine mount (1:00 and 11:00) • Low pressure recoup manifolds (from diffuser)
T12INLET TEMP SENSOR T12 SENSOR
HPT COOLING AIR THERMOCOUPLE MAIN FUEL PROBE (8) SUPPLY HOSE VSV FUEL TUBES ACTUATION LEVER
HP RECOUP FORWARD AIR MANIFOLD ENGINE MOUNT VSV ACTUATOR (2) CORE COMPARTMENT COOLING AIR
REAR ENGINE MOUNT
EEC OIL TANK
OIL SCAVENGE FILTER C2 IGNITER PLUGS VSV ACTUATOR IDG AIR/OIL HEAT EXCHANGER ELECTRICAL N1 SPEED SENSOR
IGNITION EXCITERS LOW PRESSURE TURBINE CASE COOLING
VARIABLE BYPASS VALVE (12) VARIABLE BYPASS VALVE ACTUATOR (2)
FAN FRAME
FAN ROTOR SPINNER
FAN ROTOR BLADES
NO. 1 BEARING VIBRATION SENSOR
ENGINE COMPONENT LOCATIONS B767-3S2F Page - 29
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ACC MANIFOLDS ACCESSORY LP RECOUP HEAT SHIELD AIR MANIFOLDS ACCESSORY GEARBOX TRANSFER GEARBOX
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B767-3S2F Page - 30
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GENERAL - ENGINE MOUNTS Purpose The forward and aft engine mounts transfer engine thrust and absorb vertical and side loads. The mounts allow axial and radial growth due to thermal expansion. General Component Locations The forward mount is attached to the fan frame aft inner flange and the aft mount is attached to the turbine rear frame. Inspection/check or removal/installation of either engine mount requires removal of the engine. Characteristics Forward Lower Engine Mount - This mount provides suspension of the engine at three points. The two thrust links are attached to the inner fan frame on either side of the mount assembly. The forward lower engine mount is attached to the strut by four tension bolts. Aft Lower Engine Mount - The mount lower fitting suspends the engine at two points from a double flange on the turbine rear frame. The upper fitting is attached to the strut by four bolts and barrel nuts. One point incorporates a tangential link. The aft mount transfers side, vertical and torque loads.
UPPER AFT ENGINE MOUNT TANDEM BARREL NUT (2) FAILSAFE CLEVIS ENGINE MOUNT PLATFORM AFT SHEAR PINS
YOKE
THRUST REVERSER DEFLECTION LIMITER BUMPER
FWD
LOWER AFT ENGINE MOUNT
PLATFORM LINK (2) FRAME LINK
FAN FRAME
FAN FRAME AFT INNER FLANGE TANGENTIAL LINK
TENSION BOLT (4)
LOWER FORWARD ENGINE MOUNT AFT MOUNT FORWARD MOUNT
FAN FRAME
ENGINE MOUNTS B767-3S2F Page - 31
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TURBINE REAR FRAME
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FWD
B767-3S2F Page - 32
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GENERAL - ENGINE BORESCOPE INSPECTION PORTS General Inspection of the internal parts of the engine is primarily done by means of the borescope. The engine has access for borescope inspection of each stage of the high pressure compressor, high pressure and low-pressure turbine inlets, and from ports at Stages 2 and 4 of the low pressure turbine. Additional borescope-access holes are provided in the compressor rear frame for the inspection of combustion liner and first stage turbine nozzle. A hand-operated or motor-driven system is available to facilitate borescope viewing of all high pressure rotor blades. This mounts to the accessory gearbox.
B1-10 B1-3
B1-10 B1-7
B1-10
B1-11
B4-2
B4-3
B4-1
B1-9
B1-4
B1-13
B4-2
B4-3
B1-8
B4-1 B1-13
B4-4
B1-2
B3-2
B3-1
B1-8
B1-12
B1-2
B1-6
B1-5 B5 MOTOR MOUNT (HP ROTOR BORESCOPE) B2-6
B2-1
B2-5 B2-2 B2-4 B2-3
COMBUSTION CASE LINER (AFT LOOKING IN)
ENGINE BORESCOPE INSPECTION PORTS B767-3S2F Page - 33
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B1-1
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GENERAL - ENGINE VENTS AND DRAINS
• Starter Pad • IDG Pad
Purpose Drain Mast The engine vents and drains system collects and discharges drain fluids overboard. General Description The drain system is divided into two parts. A drain module retains fluids until expelled during flight and the drain mast discharges fluid directly overboard through the drain mast. The oil tank scupper drain and combustion chamber drain are not connected to the drain module or drain mast. General Component Locations The drain module is mounted to the aft side of the engine accessory gearbox. A drain mast is attached to the fan stator case and protrudes through the engine cowling into the airstream. Drain Mast and Module The drain module is bolted on the engine accessory gearbox lower backside and is accessed by opening the thrust reverser. The drain mast is bolted to the engine fan stator case rear underside, and extends below the fan cowl. Drain Module The accessories shown in the graphic have seperate drain cavities in the drain module for storing leakage. When proper airspeed is reached the spring loaded valve inside the module opens to admit air. This air flow empties the drain cavities and discharges any accumulated fuel and oil overboard through the drain mast. The module also has push-to- open drain valves on the bottom. Each drain valve is labeled for identification. Drain valves are provided for the following components: • Hydraulic Pump Pad • Main Fuel Pump Pad • Hydro Mechanical Unit (HMU) Mount Pad
An ambient air inlet port provides air flow to the drain module. The drain lines that exit directly through the main drain are • • • • • • •
Strut Drain Left and Right Variable Stator Vane (VSV) actuator Left and Right Variable Bleed Valve (VBV) actuator Fuel Line Shroud Fuel Drain Manifold Forward Electrical Junction Box IDG Pressure Relief Valve
SCUPPER DRAIN
COMBUSTOR DRAIN LINE
OIL TANK (REF)
COMBUSTOR DRAIN VALVE (REF)
FUEL PUMP TO DRAIN MAST
HMU STARTER
DRAIN MAST
FWD
Page - 35
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FUEL MANIFOLD
B767-3S2F
DRAINS
PYLON
ENGINE VENTS AND DRAINS
FLUIDS J-BOX
SAMPLING PLUGS DRAIN MANIFOLD
OIL/HYD
IDG
FUEL AGB
FUEL DIRECT
HYDRAULIC PUMP
B767-3S2F Page - 36
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ENGINE CHANGE Engine Removal • • • • • • • • • •
Remove the fan cowl panels Open the thrust reverser doors Remove the core cowl doors Remove starter for use on engine being installed Install cover over variable bypass valve Disconnect Engine Remove the engine drain mast Install bootstrap equipment Disconnect the engine mounts Perform a general visual inspection for corrosion, powerplant strut
Engine Installation • • • • • • • • • • • • • • • • • • • • • •
Install new barrel nuts in the aft engine mount pylon fitting Prepare engine mounts for engine installation Install new serviceable mount nuts on forward engine mount Verify the Serial Number on the serviceable tag matches the Serial number on the engine data plate Provide OK to install engine Install Engine Remove cradle from engine and lower to transport stand Remove forward and aft bootstrap equipment Install the bolts on each side of the strut Install access panel for the skirt fairing Tighten the thrust links to platform attach bolts. Install the bolt and nut retainers on the forward mount Inspect mount bolt installation Install starter Drain the starter oil, check the starter magnetic chip detector and replenish the starter with oil Connect thrust reverser opening hydraulic lines Connect the strut drain line EQ Connect the drain lines for the strut raceway Install the drain mast Connect the line to the pre-cooler inlet duct Connect the hydraulic lines Install the pneumatic starter duct
• • • •
Connect the fire extinguishing discharge flex line to the tube fitting Connect pre-cooler inlet duct Connect the line to the pressure regulating valve Connect the main fuel supply line
5. FORWARD BRACKET (2 LOCATIONS)
6. UPPER AFT BRACKET (2 LOCATIONS) 7. LOWER AFT BRACKET (2 LOCATIONS)
4. AFT INBOARD ARM
8. AFT OUTBOARD ARM
3. INBOARD BRACE STRUT 2. FORWARD SUPPORT
9. DYNAMOMETER
A
1. FORWARD INBOARD ARM
SEE
E
A 11. OUTBOARD BRACE 10. AFT HOIST (2 LOCATIONS)
D
2
A
12. CABLE
13. SHEAVE (2 LOCATIONS)
E
14. DYNAMOMETER (2 LOCATIONS)
16. FORWARD OUTBOARD ARM
17. BOLTS
1
D CRADLE FWD
15. FORWARD HOIST (2 LOCATIONS) 2 18. BOLTS
OUTBD
19. BOLT
LEFT ENGINE IS SHOWN (RIGHT ENGINE IS EQUIVALENT)
22. BOLTS 20. BOLT 20A. BOLT
1
MAKE SURE THE FACE OF THE DYNAMOMETER IS AFT
2
CAUTION: _______ HOIST ASSEMBLIES MUST BE ORIENTED AT THE TOP SO THAT SLACK CHAIN WILL DESCEND FREELY BY FORCE OF GRAVITY
FWD 20C. BOLT
A
ENGINE CHANGE B767-3S2F Page - 37
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21. SKIRT FAIRING ACCESS PANEL
20B. BOLT (5 LOCATIONS)
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POWERPLANT ENGINE PRESERVATION General The GE engine must be stored and preserved against corrosion, liquids, debris and atmospheric conditions. There are three periods of preservation: • Up to 30 days • Up to 3 months • 3 months to 1 year. Preservation All engines removed from an aircraft, serviceable or unserviceable, must be preserved to the 30-day preservation procedures per the applicable Engine Maintenance Manual prior to movement into the serviceable/unserviceable engine storage areas. This preservation shall include vapor proof paper, moisture indicators and dehydrating agent even if the 30-day preservation procedures do not require it. The vapor proof paper is used to cover the intake, fan exit, and turbine exhaust. All other openings on the engine must be capped, covered, bagged, and/or protected from damage and/or contamination.
767-400 MAINTENANCE MANUAL
CF6-80C2 SERIES ENGINES
POWER PLANT - MAINTENANCE PRACTICES (PRESERVATION AND DEPRESERVATION) 1. General A. This section contains instructions for preservation and depreservation of installed power plants. Preservation consists of protecting a power plant against corrosion, liquid and debris entering the power plant, and atmospheric conditions during periods of storage, inactivity, or following an in-flight shutdown. Depreservation consists of restoring a preserved power plant to service. B. The procedure to be followed in the preservation and depreservation of an installed power plant will vary depending upon the length of inactivity, and the type of preservation used. NOTE: For engines that do not operate, refer to the preservation procedures in the GE Engine Manual. (1) The preservation procedure is based upon the following schedule: (a) Up to 30 days. (b) Up to 3 months. (c) Three months to 1 year (d) Indefinite. NOTE: There is no restriction on the number of times the preservation procedure can be renewed, as long as it is accomplished every year. C. The procedures in this section are given as a guide in deciding what precautions should be exercised to provide adequate protection from the elements during periods of inactivity. The power plant preservation schedule is a flexible program that should be implemented in a manner which suits the particular weather and storage conditions involved. A program for inactive power plants exposed to high humidity and/or large temperature changes, especially if near salt water, would require more attention to preservation needs than those engines stored in dry climates. D. The preservation program for inactive power plants must be planned in advance to implement the preservation renewal requirements, and monitored regularly to assure that the required action is implemented prior to the expiration of the preservation period.
E. The effectiveness of the preservation measures implemented should be evaluated for determining the need to extend or shorten the periods between preservation action. To be most effective, power plants in nacelles should be desiccated, and inlet and exhaust openings plugged, to help dehumidify the interior of the power plant. Humidity indicators might be helpful in monitoring moisture conditions inside the power plant even though the nacelle cannot be completely sealed from the weather. F. When desiccants are used, they must be changed on a regular basis, determined by the environmental conditions, to keep the desiccant effective. G. It is recommended that the variable bypass valve (VBV) doors be pumped closed any time the power plant is to be preserved and stored or maintenance is being performed in the area. This will avoid the possibility of foreign objects entering the core engine inlet through the VBV doors.
ACCESSORY GEARBOX (REF) HYDROMECHANICAL UNIT METERING VALVE HEAD SENSOR VSV ROD PORT VSV HEAD PORT UPPER PCB PORT VBV OPEN PORT PCR REGULATED REFERENCE PRESSURE PORT wdmt-71-00-0017
EFFECTIVITY
FWD
71-00-03
ALL H01A BOEING PROPRIETARY - Copyright (C) - Unpublished Work - See title page for details.
ENGINE PRESERVATION B767-3S2F Page - 39
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Page 202 Apr 22/07
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OIL SYSTEM - DISTRIBUTION SYSTEM OPERATION System Control The engine oil distribution system is completely automatic in operation. Pressure Oil Flow Engine oil which is stored in the oil tank flows by gravity through the supply inlet screen to the lube and scavenge pump. The pressure pump element of the lube and scavenge pump provides the motive force for lubricating and cooling the engine bearings and gears. The oil flows from this pressure pump, through the lube filter. (An oil filter service shutoff valve is provided for filter maintenance.) From the oil filter the oil flows up through a gravity loop (which keeps the oil from flowing from the tank to the bearings after engine shutdown) and out to the bearings and gears. Lubrication and Cooling The oil pressure line to the A sump distributes oil to the No. 1 (ball) bearing, Nos. 2 and 3 (roller) bearings, the accessory gear drive and bearings, and the accessory gearbox. Sump A incorporates an air/oil separator. The oil pressure line to the B and C sumps sprays oil on the No. 4 (ball), 4 (roller) and 5 (roller) bearings. Oil is sprayed on the vent tube that vents air from the B and C sumps to the A sump to reduce coking on the vent tube. The oil pressure line to the D sump sprays oil on the No. 6 (roller) bearing. Scavenge Oil Flow Oil from the A sump drains down the radial drive shaft housing into the transfer gearbox where it is scavenged. A slinger-type disk pumps in the A and D sumps provide positive sump draining for high altitude operation or airplane maneuvers when scavenge would otherwise be difficult. The oil from the sumps and the gearboxes returns to the Lube and Scavenge Pump via inlet screens to the five scavenge pump elements. All scavenge oil flow from the five scavenge pump elements is combined within the pump gallery to be discharged at one common port .
From the lube and scavenge pump the scavenge oil flows under pressure past the magnetic chip detector and then through the servo fuel heater and the fuel/ oil heat exchanger. The scavenge oil flow is then cleaned by the scavenge oil filter as it returns to the oil tank. Note:
The lubrication system is fully operational only when the engine is running. It is not fully operational when the engine is motoring or wind milling. Motoring and wind milling operations do not provide adequate sump seal pressurization nor sufficient scavenge flows. Consequently, increased apparent oil consumption rates and abnormal oil hiding occur.
DEAERATOR
SCAVENGE OIL FILTER
OIL QTY XMTR
B
OVER FILL
SIGHT GLASS
PRESS RELIEF VALVE
OIL FILTER
~P
D SUMP
B/C SUMP
A SUMP SLINGER DISK PUMP
OIL LINE PRESS SUPPLY SCAVENGE PUMP IN SCAVENGE PUMP OUT VENT OIL JET BALL BRG ROLLER BRG MAG DET OIL STRAINER DRAIN PLUG
FLAME ARRESTOR ENG OIL PRESS XMTR TRANSFER GEAR BOX
ENG LOW OIL PRESS SWITCH PRESS PUMP
ACCESSORY GEAR BOX
ANTI-STATIC LEAK VLV
LUBE AND SCAVENGE PUMP FUEL/OIL SERVO FUEL EXCHANGER HEATER
PR
OIL TEMP SENSOR
MAGNETIC CHIP DETECTOR
OIL DISTRIBUTION SYSTEM OPERATION B767-3S2F Page - 41
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OIL SYSTEM - OIL STORAGE SYSTEM Storage System Components The oil storage system consists of the following components: • • • •
Oil Tank Oil Tank Filler Cap Oil Tank Pressurizing Valve Oil Tank Pressure Relief Valve
Oil Tank The oil tank provides storage for the engine oil. It is located on the right side of the fan case. Access is gained by opening the right fan cowl panel. It is constructed of aluminum and may have an external coating of a silicone rubber compound for insulation. A plug for oil draining is provided on the bottom of the oil tank. Oil Tank Filler Cap The oil tank filler cap allows manual filling of the oil tank and seals the opening of the fill port. The filler cap is located on the upper right side of the oil tank. access for servicing may be gained by opening the oil tank access door located on the right fan cowl panel or by opening the right fan cowl panel. Oil Tank Pressurizing Valve The oil tank pressurizing valve maintains tank internal pressure. The pressurizing valve is located on top of the oil tank. Access is gained by opening the right fan cowl panel. The oil tank is pressurized by the returning air-oil stream. The oil tank pressurizing valve vents air into the A sump at 7-11 psi above the transfer gearbox vent pressure. Pressure Relief Valve The pressure relief valve is a back-up safety valve that relieves tank pressure. at 27 psi venting to ambient air preventing tank rupture. The relief valve is located below the fill port scupper. Access is gained by opening the right fan cowl panel.
CAUTION: DO NOT OVERFILL. IF ENGINE HAS BEEN MOTORED WITHOUT SUBSEQUENT OPERATION FOR SCAVENGING, OIL LEVEL WILL BE APPROXIMATELY TWO QUARTS (TWO LITERS) LOW.
SCAVENGE RETURN TUBE
OIL TANK FILLER CAP
VENT TUBE PRESSURIZING VALVE
FILLER CAP SCUPPER DRAIN TUBE
ACCESS DOOR
ENGINE OIL TANK
ENGINE OIL TANK OVERFILL PORT SCUPPER DRAIN
DRAIN PLUG
OIL SUPPLY TUBE
SIGHT GLASS
OIL STORAGE B767-3S2F Page - 43
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PRESSURE FILL PORT
PRESSURE RELIEF VALVE
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OIL SYSTEM - LUBE AND SCAVENGE PUMP Purpose The Lube and scavenge pump provides the motive force for the lubricating oil. Location and Access The lube and scavenge pump is mounted on the forward side of the accessory gearbox. It is accessible when the Thrust Reversers are open. Characteristics The lube and scavenge pump contains one pressure pump element and five scavenge pump elements. In the pump housing are two rows of vane type positive displacement pumps. Each row contains three pumping elements. The difference between the pumping elements is capacity which is determined by the diameter and length of each. No regulation of oil pressure is provided within the oil pump. Power The lube and scavenge pump is spline shaft driven by the accessory gearbox.
ACCESSORY GEARBOX SCAVENGE INLET D SUMP SCREEN SCAVENGE INLET SCREEN
FROM B AND C SUMP
FROM TRANSFER GEARBOX
FROM OIL TANK
TO ENGINE BEARINGS AND GEARBOXES ANTI STATIC LEAK VALVE
LUBE SUPPLY INLET SCREEN
FWD
PRESSURE PUMP
ACCESSORY GEAR BOX C SUMP SCAVENGE INLET SCREEN
B SUMP SCAVENGE INLET SCREEN
A SUMP AND TRANSFER GEARBOX SCAVENGE INLET SCREEN
FROM D SUMP SCAVENGE OIL TO OIL TANK OIL LINES
FWD
PRESSURE
SCAVENGE PUMP OUT
SUPPLY
MAG DET
SCAVENGE PUMP IN
OIL STRAINER DRAIN PLUG
LUBE AND SCAVANGE PUMP B767-3S2F Page - 45
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OIL SYSTEM - MAGNETIC CHIP DETECTORS Magnetic Chip Detectors The magnetic chip detectors attract metallic particles carried in the scavenge oil. One is provided for each scavenge pump as well as a master chip detector for all scavenge oil on return. The master chip detector is located in the scavenge discharge flow tubing adjacent to the drain module. The individual scavenge pump chip detectors are located on the inlet side of the respective scavenge pump, and are saftied to the pump with safety wire. Access is gained by opening the integrated drive generator service door or by opening the thrust reversers. Characteristics The magnetic chip detector is a permanent magnet probe. An internal check valve permits removal of the chip detector probe for inspection without draining the oil system. CAUTION: WHEN REMOVING CHIP DETECTOR ENSURE A SERVICABLE “O” RING IS INSTALLED UPON INSTALLATION.
DRAIN MODULE OIL TUBE
OIL FLOW FROM SCAVENGE PUMPS HOUSING
MAGNETIC CHIP DETECTOR
MAGNETIC CHIP DETECTORS B767-3S2F Page - 47
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OIL SYSTEM - SCAVENGE OIL FILTER AND HEAT EXCHANGERS Scavenge Oil Filter The scavenge oil filter, in conjunction with the lube filter and the supply and scavenge inlet screens, clean contaminants from the oil. Characteristics The scavenge oil filter is of the replaceable element type. A filter relief valve is provided that begins bypassing oil at approximately 40 psid for a partially clogged filter. At 60 psid the relief valve is fully open. The scavenge oil filter is located below the oil tank on the right side of the fan case. Access is gained by opening the right fan cowl panel. Fuel Oil Heat Exchanger The fuel/oil heat exchanger dissipates oil heat and heats the fuel. Characteristics The fuel/oil heat exchanger consists of a multi-tube core, mounted in a cylindrical housing that contains two inlet ports and two outlet ports. One set of ports is used for fuel passage through the tubes of the heat exchanger core. The other set of ports allows passage of oil around the core tubes within the housing. All engine fuel passes through the heat exchanger since there is no provision for bypass. A pressure relief valve permits scavenge oil to bypass the core tubes at engine start up during cold weather. The fuel/oil heat exchanger is bolted to the fuel pump on the bottom right side of the engine. It is accessible when the thrust reversers are open. Servo Fuel Heater The servo fuel heater is used for additional heating of the fuel specifically used for hydraulic movement of components.
Characteristics The servo fuel heater consists of a multi-tube core, mounted in a cylindrical housing that contains two inlet ports and two outlet ports. One set of ports is used for fuel passage through the tubes of the heater core. The other set of ports allows passage of oil around the core tubes within the housing. The servo fuel heater is located on the right side of the engine at the 5:00 position. It is accessible when the right thrust reverser is open.
OIL TANK
OUTLET PORT
OUTLET PORT
SCAVENGE OIL FILTER
FUEL/OIL HEAT EXCHANGER
FAN CASE IN
IN
FUEL FLOW
PACKING
OUT
FILTER ELEMENT
OIL FLOW OUT FWD
OIL SCAVENGE FILTER BOWL
PRESSURE RELIEF VALVE
SCAVANGE OIL FILTER AND HEAT EXCHANGER B767-3S2F Page - 49
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FILTER HEAD
INLET PORT
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OIL SYSTEM - OIL DISTRIBUTION SYSTEM Purpose The oil distribution system provides supply and scavenge force for lubricating the engine bearings and gearboxes, for cooling the oil, and for cleaning any contaminants from the oil. General Component Locations The system component can be located inside the right thrust reverser and fan cowls. System components are: • • • • • •
Lube and Scavenge Pump Scavenge Oil Filter Engine Lube Filter Fuel/Oil Heat Exchanger Servo Fuel Heater Magnetic Chip Detectors
General Operation All functions of the oil distribution system are completely automatic in operation.
MAGNETIC CHIP DETECTOR
SERVO FUEL HEATER DRAIN MODULE SCAVENGE OIL FILTER ACCESS GEARBOX
LUBE AND SCAVENGE PUMP
FUEL/OIL HEAT EXCHANGER
OIL DISRIBUTION SYSTEM B767-3S2F Page - 51
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FUEL PUMP
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OIL SYSTEM - OIL INDICATING SYSTEM General The oil indicating system includes: • • • • •
oil quantity oil temperature oil pressure low oil pressure oil filter bypass indicating
Oil indication appears on EICAS. A L(R) ENG OIL PRESS light for each engine is located below the Standby Engine Indicator. Indications All oil pressure indications are visible on the Secondary Engine display and the “PERF / APU” page. The engine oil temperature indication is provided to EICAS from the EEC. Also, the following messages are displayed on the primary engine display: • L / R ENG OIL PRESS (C) • L / R OIL FILTER (C) In the case of the “Low Oil Press” indication two engine discrete lights are located directly under the SEI. These lights indicate “L / R ENG OIL PRESS”. The lights are normally on with the engines shut down and input for these comes directly from the low oil pressure switch on the engine.
N1 AUTO ON EGT N2
L ENG OIL PRESS a
LOW OIL PRESSURE SWITCH OIL PRESSURE TRANSMITTER
L (R) ENG 0IL PRESS L (R) OIL FILTER
OIL TEMPERATURE SENSOR
PRIMARY ENGINE DISPLAY
EEC
OIL QUANTITY TRANSMITTER
EICAS COMPUTER
OIL INDICATING SYSTEM Page - 53
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70
35
OIL
PRESS
105
70
OIL
TEMP
18
3
OIL
QTY 70 105 18
OIL FILTER DIFF PRESSURE
B767-3S2F
R ENG OIL PRESS a
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SECONDARY ENGINE DISPLAY
OIL PRESS 35 OIL TEMP 70 OIL QTY 03
PERF/APU PAGE
B767-3S2F Page - 54
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OIL SYSTEM - OIL INDICATION OPERATION Oil Quantity The oil quantity transmitter provides a reference signal to the EICAS computers for determining the level of oil in the tank. The oil quantity transmitter is mounted into the top of the rear half of the oil tank. Access is gained by opening right the fan cowl. Oil Quantity appears on the EICAS Secondary Engine Display and on the PERF/APU page. The oil quantity transmitter contains a sealed liquid-level sensing unit. The sensing unit is a hollow tube containing magnetic reed switches and a resistor network, a cylindrical float houses a permanent magnet. The indicator unit is line replaceable. Oil Pressure Transmitter Oil pressure appears on the EICAS Secondary Engine Display and on the PERF/APU page. The oil pressure transmitter senses the differential pressure between the oil supply manifold and the accessory gearbox vent. The oil pressure transmitter is mounted on a bracket adjacent to the lube filter. Access is gained by opening the right thrust reverser. Oil Pressure Limits The lower red line limit for oil pressure is 10 psid. The yellow band upper limit changes between idle and full power as a linear function of N2. The yellow band upper limit is 13 psid when the engine is at low idle (60% N2). At full power (110% N2), the yellow band upper limit is 34 psid. Low Oil Pressure Switch The low oil pressure switch senses the differential pressure between the oil supply manifold and the accessory gearbox vent. It is bracket-mounted adjacent to the lube filter. Access is gained by opening the thrust reverser. The switch contacts are normally closed. The switch opens at 15 psid and closes at 10 psid. When the oil pressure is low, the switch illuminates the low oil pressure warning light and the message L(R) ENG OIL PRESS appears on EICAS.
Oil Temperature Sensor The oil temperature sensor is a thermocouple probe which sends a digital signal to EICAS. Oil temperature is indicated on the EICAS secondary engine display and on the PERF/APU page. 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. The operational range of the TEO sensor input to the EEC is from -81 to 352 degrees F(-63 to 178 degrees C). The red line limit is 347 degrees F (175 degrees C). The yellow band range is from 320 degrees F(160 degrees C) to the red line limit. Oil Filter Differential Pressure Switch The oil filter differential pressure switch is a diaphragm-controlled snap-action normally opens the switch that closes when the differential pressure across the scavenge filter element is 25 - 33 psid. The switch configuration is normally open. 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. The EICAS message will extinguish when the switch opens at 25 psid. Or less.
SCAVENGE OIL FILTER FROM HEAT EXCHANGERS
15 PSID
A
L ENG OIL PRESS a 117.5
960
AUTO ON
>33 PSID
L ENG OIL PRESS a
LUBE AND SCAVENGE PUMP PRESSURE OUTPUT
1.
L ENG OIL PRESS (C)
ENG OIL PRESSURE EICAS
E E C
RESISTOR SWITCH NETWORK
REF PWR
105
EEC
L EICAS
R EICAS 22 SWITCHES.
2 OR 3 SWITCHES MAGNETICALLY CLOSED AT ANY LEVEL
OIL INDICATING SYSTEM B767-3S2F
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SECONDARY ENGINE DISPLAY
OIL
TEMP
QUARTS
18 OIL
3 QTY
LITERS
PERF/APU
REF PWR DC REF
1
PRESS
CHAN B
1
OIL QUANTITY TRANSMITTER
35
DC REF
FROM SCAVENGE OIL TEMP PUMPS SENSOR
EMPTY
70 OIL
CHAN A
OIL PRESSURE TRANSMITTER
MAGNET
1.
PRIMARY ENGINE DISPLAY
TO FUEL/OIL HEAT DUAL ELEMENTS EXCHANGERS
P11
Page - 55
R ENG OIL PRESS a
L OIL FILTER (C)
28V DC L BUS
OIL PRESS LINE
112.5
N2
OIL PRESSURE LIGHT
LOW OIL PRESSURE SWITCH
OIL FILTER DIFFERENTIAL ACCESSORY PRESSURE SWITCH GEARBOX (VENT)
960
EGT
112.5
86%
N1 ACT
>150 RPM/SEC
5% N2 ACT
SAME AS CH A
ACCELERATION COMMAND DETECTOR EEC
FAN AIR OPEN
CCCV SOL
INTERNAL ENG COOLING AIR FLOW TO HPT SECOND STAGE NOZZLES AND BLADES
11TH STAGE AIR
UNCONTROLLED
CCCV CONTROL B767-3S2F Page - 83
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ENGINE AIR SYSTEM - CCCV SYSTEM General The core compartment cooling system supplies controlled cooling air for the core-mounted engine accessories. The system decreases the core cooling at low power and high altitudes to conserve primary air. The system has one Core Compartment Cooling Valve (CCCV). The valve is controlled by the CCCV solenoid. The EEC controls the solenoid. Core Compartment Cooling Valve (CCCV) The core compartment receives fan air for cooling through the CCCV and manifold. The valve is located at the 10:00 position on the HPC case. The butterfly-type valve is spring-loaded 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 indicates valve position. The manifold sends airflow to these items: • • • •
HPC case IDG Hydraulic pump Fuel pump
CCCV Solenoid The CCCV solenoid controls the flow of eleventh-stage air. The solenoid valve is spring-loaded closed. The eleventh stage air pressure comes from the supply duct on the left side of the engine. When the solenoid is energized, the eleventh-stage air pressure is directed to the CCCV to close it. MAINTENANCE TIP To remove the valve, move the butterfly to the closed position against spring pressure. The butterfly valve shaft is attached to the valve position indicator with a roll pin. The valve position indicator has a hexagonal nut casting that can be moved with a 7/16-inch wrench. CAUTION: IF YOU USE TOO MUCH TORQUE DURING MANUAL CLOSING OF THE VALVE, THE ROLL PIN WILL SHEAR. THIS CAUSES
THE BUTTERFLY VALVE TO STAY IN THE OPEN POSITION AND YOU CAN NOT REMOVE THE VALVE WITHOUT REMOVAL OF ADDITIONAL DUCTING.
MANUAL/LOCK OPEN SCREW/PIN STOWAGE
CLOSED
ELECTRICAL CONNECTOR 11TH STAGE AIR
SOLENOID
OPEN
VALVE POSITION INDICATOR TOP VIEW
CHANNEL A 16V DC CHANNEL B
CORE COMPARTMENT COOLING MANIFOLD
EEC FLOW ARROW BUTTERFLY VALVE FAN AIR DUCT
CORE COMPARTMENT COOLING VALVE (CCCV)
CCCV SYSTEM B767-3S2F Page - 85
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ENGINE AIR - TURBINE CASE COOLING Introduction The turbine case cooling (active clearance control) system uses separate manifolds to cool the LPT and HPT cases. The HPTC valve controls the fan air to the HPT manifold. There is no valve for the LPTC manifold. The LPTC and HPTC manifolds send fan air onto their respective turbine cases. This decreases case expansion which decreases turbine blade tip-to-case clearance and increases turbine efficiency. Description The HPTC valve is located on the right side of the engine at the 1:00 position near the eleventh-stage bleed manifold. HPTC Valve A hydraulic piston actuator controls the butterfly-type HPTC valve. Hydraulic fluid pressures received from Electro-Hydraulic Servo Valve (EHSV) in the HydroMechanical Unit (HMU) controls the modulation of the valve. The EEC controls the EHSV. The valve assembly has two Linear Variable Differential Transformers (LVDTs) which supply valve position signals to the 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. The valve is commanded open when the pressure altitude is above 15,000 feet and N2 speed is between 82 and 98 percent. Operation These are software components in the EEC channel processors: • • • •
Turbine growth calculators HPTC command calculators Demand calculators Valve drivers
The growth calculators receive multiple engine sensor inputs and insure the size of the inner diameter of the turbine case is equal to the size of the outer diameter of the rotor plus the desired clearance.
HPTC VALVE
HPTC VALVE POSITION FEEDBACK
N1 ACT
FEEDBACK
HPT DIMENSIONAL
N2 ACT
CALCULATOR
PT
SIZE ERROR
TAT
HPTC COMMAND CALCULATOR
HPT CMD
HPTC DEMAND CALCULATOR
HPT DMD
HPTC VALVE DRIVE
HPTC EHSV
SERVO REGULATOR
PO PS3 HMU
T25 T3 T49 ACTIVE CHANNEL EEC
TURBINE CASE COOLING CONTROL B767-3S2F Page - 87
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SERVO FUEL IN
B767-3S2F Page - 88
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ENGINE AIR CONTROL - TURBINE CASE COOLING (TCC OR ACC) Description The turbine case cooling system uses separate manifold to cool the LPT and HPT cases. The fan air to the HPT manifold is controlled by the High Pressure Turbine Cooling Valve (HPTCV). Then LPTC and the HPTC manifolds encircle and direct fan air onto their respective turbine cases. This reduces case expansion, thus minimizing turbine blade tip to case clearance which increases turbine efficiency. The HPTCV 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. HPTCV The HPTCV is a butterfly type valve controlled by a hydraulic piston actuator. Modulation of the valve is operated by a hydraulic fluid pressure received from an EHSV on the Hydro Mechanical 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 the EEC. There is an electrical connector for each LVDT. One LVDT delivers feed back to channel A and the other to channel B of the EEC.
FAN AIR SUPPLY DUCT
LPTC MANIFOLD
HPTC MANIFOLD
HPTC
VALVE
FAN AIR SUPPLY DUCT
ACTUATOR
LVDT
ROD END HEAD END
REF
CH A HPTC SERVO VALVE (EHSV) CH B
PRESS EHSV PRESS
BUTTERFLY VALVE
FLOW ARROW EEC
HMU VALVE (TYPICAL)
TURBINE CASE COOLING (ACTIVE CLEARANCE CONTROL) B767-3S2F Page - 89
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ENGINE AIR - INDICATIONS General Position indications show on the EPCS page for these engine air system components: • Variable Stator Vane (VSV) actuators • Variable Bypass Valve (VBV) actuators • High Pressure Turbine Cooling (HPTC) valve These parameter values show on the EPCS page for the temperatures and pressures for control of engine air system components: • • • •
Ambient (static) pressure (P0) HPC discharge (burner) static pressure (PS3) HPC inlet temperature (T2.5) HPC discharge (burner) temperature (T3)
The indications are in percent of maximum angle, with 0 percent equal to fullyclosed positions and 100 percent equal to fully-open. The ranges for the indications are from -5.0 percent to 105.0 percent. The P0 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 160C, and the T3 indication range is from -55 to 650C. A box surrounds the EEC channel that is in control.
EPCS A
-5.0 -5.0 -127.5 -80 -1.5 -5 -5.0 -5.0 -55 -55 -5 1
DISPLAY VALUE LIMITS
2
TYPICAL IDLE VALUES
3
TYPICAL CRUISE VALUES
B
A
105.0 105.0 127.5 90 20.0 105 105.0 105.0 160 650 600
1
VSV VBV TRA T 12 P0 HPTC T/R L T/R R T 2.5 T3 PS3
1
ENGINE AIR SYSTEM EICAS INDICATIONS Page - 91
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25.3 58.5 33.9 15 14.7 0 0.0 0.0 18 172 48 2
EPCS PAGE
B767-3S2F
B
84.3 0 71.1 15 14.7 35 0.0 0.0 90 504 381 3
B767-3S2F Page - 92
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ENGINE AIR - ENGINE AIR SYSTEM - OPERATION Variable Stator Vanes The VSVs move from fully closed during starting to fully open at takeoff power. The modulation schedule changes during reverse thrust operation. The VSVs fail-safe closed. Variable Bypass Valves The VBVs move from fully open during starting to fully closed at takeoff power. The modulation schedule changes during rapid deceleration and reverse thrust operation. The VBVs fail-safe open. Core Compartment Cooling Valves The CCCV is closed at stabilized cruise power when the aircraft is above 17,000 feet altitude and the EGT is less than 699C. Cooling airflow to engine accessories is reduced when the CCCV is closed. The CCCV is fail-safe open. HPTC Valve The HPTC valve opens at cruise power settings when the aircraft is above 17,000 feet altitude and N2 is between 82 percent and 98 percent. Turbine case cooling airflow is increased when the valve is open. The HPTC valve is fail-safe closed.
NAME OF SUBSYSTEM
ENGINE SHUT DOWN
IDLE
TAKEOFF POWER
CRUISE
RAPID DECEL
FAIL/ SAFE
CORE COMPARTMENT COOLING VALVE
REV
N/A 1
1
1
HPTC VALVE
N/A
VARIABLE STATOR VANES (VSV) 2 VARIABLE BYPASS VALVES (VBV) 3
= MODULATING = OPEN = CLOSED = REDUCED FLOW
1
ABOVE 17,000 FT, N2 STABILIZED, EGT LESS THAN 699C
2
MOVE 4 DEGREES TOWARDS CLOSE
3
OPEN ADDITIONAL 30 IN2
ENGINE AIR SYSTEM - OPERATION B767-3S2F Page - 93
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3
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ENGINE INDICATING SYSTEM - SYSTEM OVERVIEW General Engine indicating systems include • • • • • •
Power Indication Vibration Monitoring Temperature Indication Power Management Control Monitoring Oils System Indication Fuel System Indication
Power Indication The primary power indication is the Low Pressure Rotor Speed, or N1, given in percent rpm. The N1 Rotor Speed is measured by an Electromagnetic Sensor and a 38 Tooth Rotor in the "A" Sump of the engine. An Electromagnetic Pulse is generated in the Sensor Coils each time a Tooth passes. The pulses per unit of time are measured by the EICAS Computers, Standby Engine Indicator, EEC, or by the Fan Trim Box as appropriate, and converted to an N1 rpm signal. The signal is presented on the upper EICAS display. The signal is displayed digitally on the Standby Engine Indicator. The signal is used by the EEC for Computations and trimming. The N2 Rotor Speed is provided by a seperate sensor mounted to the front of the accessory gear box. The N2 sensor generates a Frequency that is proportional to N2 Rotor Speed. The EICAS Computers and Standby Engine Indicators convert the Frequency to a N2 rpm Display. The N2 is presented on the Lower EICAS Display Unit when the "ENGINE" EICAS switch is selected. The N2 is displayed Digitally on the Standby Engine Indicator also.
Temperature Indication (EGT) An Exhaust Gas Temperature indication is used to monitor the Engine Temperature. Thermocouple Probes are located between the High and Low Pressure Turbines. The EGT system utilizes Eight (8) Chromel-Alumel Thermocouple Probes installed on the Low Pressure Turbine forward case (Station T49). The probes are electrically connected in parallel to provide a voltage to the EICAS Computers that is proportional to Exhaust Gas Temperature. The EGT is displayed on the Upper EICAS Display. EGT is also displayed on the Standby Engine Indicator. Propulsion Interface Monitoring Unit (PIMU) The EEC Micro-Processors are both monitored by a Propulsion Interface Monitoring Unit (PIMU) located in the Main Equipment Center. Indication that an EEC fault has been stored in the monitor is provided by an EICAS display of a "PIMU" Maintenance Message. Oil Indication systems Oil systems report information that includes: • • • • •
Oil Pressure Low Oil Pressure Oil Filter DP Indication Oil Temperature Oil Quantity
These indications are reported to EICAS as wel as the SEI for reporting, and fault annunciation in the cockpit. Fuel Indication Systems
Airborne Vibration Monitoring (AVM) Two Sensor Probes, employing Piezoelectric Crystals to sense vibration of the rotors, are utilized to monitor the engine vibration. A Vibration Monitor unit in the Main Equipment Center prepares the sensor signals for the EICAS display.
The fuel indication system reports inter-stage fuel pressure and fuel flow to the EICAS systems as well as the FMS. This is used by the FMS to calculate fuel economy along with the software profile loaded. Also, fuel differential pressure (DP) is measured across the Main fuel filter. This is reported to EICAS if this pressure differential becomes too great.
EGT PROBE (8)
N1 SPEED SENSOR CRF ACCELEROMETER
N2 SPEED SENSOR N2 SPEED CARD
ALTERNATE ACCELEROMETER
P50 PIMU (E1/E2)
NO.1 BRG ACCELEROMETER EEC
AVM
SIGNAL CONDITIONER EICAS
TAT +13c ENG 2 FIRE L GEN OFF PARKING BRAKE
PERF/APU
D-TO +15c 65
577
81.1
81.1
756
756 N1
CABIN CALL GROUND CALL
(E8)
120
577
OIL PRESS OIL TEMP
65 N2 120
6.4
6.4
15 OIL QTY 15 EGT
1.2 BB
VIB
70 OIL PRESS 70 105 OIL TEMP 105 18 OIL QTY 18 VIB FAN LPT N2 BB
1.2 0.9 1.1 1.2
1.2 0.9 1.1 1.2
FF APU: EGT 640 RPM 99
1.2 N2
GROSS WT 187.6 CAS 245 TAT +12.0 MACH 0.615 ALT 21030 85.0 N 1 MAX 81.2 N 1 CMD 81.2 N 1 ACT 141.7 TRA SEL 625 EGT N2 67.7 12.312 FF FP 86 40 DUCT PRESS
104.0 81.2 81.2 141.7 625 67.7 12.312 86 40
APU OIL QTY
PRIMARY ENGINE DISPLAY
SECONDARY ENGINE DISPLAY
PERFORMANCE/APU PAGE
INDICATING SYSTEM OVERVIEW B767-3S2F Page - 95
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ENG EXCD
R EGT REDLINE 967 965 955 945 935
:05.4 :12.3 :16.8 :19.3
ENGINE EXCEEDANCE PAGE
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ENGINE INDICATION SYSTEM - STANDBY ENGINE INDICATOR (SEI) Purpose The SEI provides backup N1, EGT and N2 indications when EICAS is unpowered, or otherwise not displaying the primary engine parameters. Features The SEI utilizes LEDs for its displays. Six displays show N1, EGT, and N2 for both engines. The unit has its own power supply and circuitry. A test switch is built in to allow testing the SEI for correct operation. The SEI indicates malfunctions on both N1 displays. A two-position switch on the face of the unit allows either AUTO or ON to be selected. In AUTO the SEI display is inhibited if EICAS primary engine parameters are available. Should both EICAS computers or both EICAS displays become inoperative, the SEI will automatically begin displaying it’s parameters if the engine is operating. The SEI display is continuous in the ON position. Interfaces The SEI receives analog input signals from the EEC on the FADEC engine. These indications are only available when the EEC is powered. Note:
The SEI as delivered from the supplier is adaptable to different model engines. The words FAIL NO LIMIT appear on the face of the indicator. The correct placard for the GE CF6-80C2F engine must be removed from the old SEI and installed on the new unit before the unit is installed in the panel.
FAIL NO LIM
EPR
FAIL NO LIM
SUPPLIER PLACARD (2)
N1
FAIL NO LIM
EGT
FAIL NO LIM
REMOVE COVERPLATE (2)
N2
AUTO/ON SWITCH
AUTO ON
TEST SWITCH
AS DELIVERED BY SUPPLIER
ADD OPERATIONAL PLACARDS (2)
EPR
117.4
N1
117.4
960
EGT
960
112.5
N2
112.5
OPERATIONAL PLACARDS
AUTO ON
AS INSTALLED ON AIRPLANE
1
STANDBY ENGINE INDICATOR (SEI) B767-3S2F Page - 97
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General
The three coil-induced speed signals are sent through two separate electrical connectors. One coil output goes through one connector to EEC channel A. The other two coil outputs go through the second electrical connector - one output to EEC channel B, and the other output to EICAS and the AVM. All three outputs are identical.
There are two engine tachometer indications. The low pressure shaft speed is called N1. The high pressure shaft speed is called N2. N1 is the primary thrust indication. An N1 speed sensor on the fan case provides the output signals. The signal is sent to the EEC, and the Airborne Vibration Monitor (AVM). The EEC forwards the information in digital format to EICAS and the SEI.
The output of the N1 sensor is also used during the fan trim balance procedure. One of the ferromagnetic teeth provided on the sensing wheel is taller than the rest, and the pulse it produces is stronger. This stronger pulse is generated once for every complete revolution of the fan shaft, and is used to track balancing errors in the fan assembly.
ENGINE INDICATING SYSTEM - ENGINE TACHOMETER SYSTEM
N2 is the secondary thrust indication. The EEC N2 speed sensor provides an N2 signal to the EEC, N2 discrete’s printed card and AVM. The EEC forwards the information in digital format to EICAS and the SEI. Exhaust Gas Temperature (EGT) Indicating System N1 SENSOR The N1 fan shaft speed sensor is mounted on the fan frame at the 2:00 position, just aft of the No. 3 strut. The N1 sensor is a magnetic speed pickup with three electrically-isolated coils located in the sensor tip. The sensor has a stainless steel housing and a mounting flange with two bolts holes. The sensor assembly is about 20 inches long and the housing is 3/4 inch in diameter. The engine has a support tube inside the No. 3 strut and a titanium receiver to hold the sensor in place. The mounting flange spring holds the sensor tip snug against the titanium receiver to prevent vibration. The titanium receiver also protects the tip from sump oil. There is a rubber bushing at the sensor housing mid-pint to prevent housing vibration. When installed, the sensor tip is in close proximity (0.10 inch nominal) to a 38tooth ferromagnetic wheel. The wheel is pressed onto the forward fan shaft in front of the No. 2 bearing inner race. As the fan shaft rotates, each tooth passes the sensor which induces a pulse in each of the three sensor coils. Thirty-eight pulses are generated during each complete revolution of the fan shaft. The pulse frequency is directly proportional to the fan shaft speed. Access to the sensor is through the right thrust reverser half. Access to the wheel requires major engine disassembly.
N2 Core Shaft Speed Sensor The N2 core shaft speed sensor has a permanent magnet and three electricallyisolated coils located in the sensor tip. The sensor has a mounting flange with two bolt holes. The sensor assembly is mounted on the forward right side of the accessory gearbox, inboard of the hydro-mechanical unit (HMU). The three coil-induced speed signals are sent through two separate electrical connectors. One coil signal goes through one connector to EEC channel A and the other two coil signals go through the second electrical connector; one to EEC channel B and the other to EICAS, AVM and the N2 speed card. The electrical outputs are AC signals whose frequency is directly proportional to core speed. The signals are generated by three electrically isolated coils located just behind a permanent magnet installed at the sensing tip of the probe. When the probe is inserted through the gearbox wall, the sensing tip is brought within close clearance (.037 inch nominal) of 12 ferromagnetic lugs installed on the forward face of an idler gear that sets between the starter drive gear and the main fuel pump drive gear. As each lug passes the tip of the sensor, it induces a voltage into each of the three coils. The starter gear is driven directly by the horizontal drive shaft, and the idler gear is driven by the starter gear.
N1 SPEED SENSOR
FAN 3 STRUT
N2 SPEED SENSOR
ACCESSORY GEARBOX (FWD SIDE)
FWD
ENGINE TACHOMETER SENSORS B767-3S2F Page - 99
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ENGINE INDICATION SYSTEM - 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. This same information can be seen on the ‘PERF/APU maintenance page. 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. This excessive speed information is also recorded on the engine exceedance page. 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 - ENGINE SECONDARY DISPLAY Actual N2 for each engine appears on the EICAS secondary engine display as a digital readout and a pointer on a round analog scale. This same information can be seen on the ‘PERF/APU maintenance page.
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 after the exceedence event has passed. This excessive speed information is also recorded on the engine exceedence page. A magenta fuel on command line appears when the engines are shut down. The value is set at 15 percent N2 on the ground and 10 percent N2 in flight. This is minimum engine speed indication for fuel command on. The analog speed information given to EICAS is compared with the N2 digital information. Should the analog signal be 40% or less and the digital signal be greater than idle for 10 seconds, the Status/Maintenance message “L/R Eng Analog” will be displayed on the EICAS Status / Maintenance page. This indication alerts maintenance to the loss of N2 speed information to the AVM and N2 Speed Card. 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 valves 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.
PERF/APU GROSS WT187.6 CAS MACH
70 105 18
OIL PRESS OIL TEMP OIL QTY
35 70 12
95.2 95.2 54.9 -21.54 528 104.2 12.436 86 40 320 -19.1
VIB
1.2 0.9 1.1 1.2
N1, N2
FAN LPT N2 BB
0.3 2.2 0.9 2.3
APU: APU:
CHANNEL A
EGT RPM
SEI
CHANNEL A CHANNEL B
245 0.615
640 87
TAT ALT N1 CMD MAX ACT TRA SEL EGT N2 FF FP DUCT PR BURN PR T/R
AUTO EVENT
APU OIL QTY
+12 21030
0.0 0.0 26.1 141.75 825 23.4 15.312 84 40 390 120.5
R EGT RED
N1, N2 AENG EXCD
CHANNEL B
121.7 903
EEC
:06 :12
N1 RED EGT START EGT RED N2 RED
MAX
903 900 :02.7 885 :03.5 870 :04.4 855 :05.2 840 :06.3 825 :07.6 810 :08.5 795 :09.1 780 :10.4 765 :11.3 750 :12.2
N1 SPEED SENSOR
114.9
R EGT AMBER MAX
955 1:09.5 957 945 1:11.2 935 1:13.3 925 1:15.7
N1 EICAS AIRBORNE VIBRATION MONITOR
CHANNEL A
FROM OTHER ENGINE
N2
DIGITAL N2 > IDLE 83%
1
28-25
72%
A
OPENS THE APU ISOLATION VALVE AND STARTS THE DC PUMP WHEN AIRBORNE WHEN L ENGINES N2 < 72% AND IS ON SUCTION FLOW
1
28-22
50%
B
SHUTS OFF THE OVERRIDE PUMP WHEN THE RESPECTIVE ENGINE N2 < 50% (K-2)
1
73-21
50%
B
INHIBITS THE "EEC INOP" AMBER LIGHT WITH THE RESPECTIVE ENGINE N2 < 50% (K4)
1
30-31
50%
B
PROVIDES LOW HEAT MODE ON PITOT-STATIC PROBES ON THE GROUND WHEN EITHER ENGINE N2 > 50% (K2)
1
80-11
50%
C
CLOSES RESPECTIVE ENGINE START VALVE BY DEENERGIZING THE START SWITCH SOLENOID AND START RELAY WHEN N2 > 50% (K1)
2
24-51
50%
D
(OPTIONAL) SHEDS A PORTION OF THE GROUND SERVICE BUS WHEN THE R UTILITY BUS IS UNPOWERED AND EITHER ENGINE N2 < 50% (AIRBORNE ONLY) (K9)
2
21-51
50%
D
INHIBITS HIGH FLOW SCHEDULE FOR OPPOSITE COOLING PACK WITH EITHER ENGINE N2 < 50% (AIRBORNE ONLY)
2
21-58
50%
D
(K6) PROVIDES "INBOARD OPEN LOOP" FOR EQUIPMENT COOLING ON THE GROUND WITH BOTH ENGINES N2 > 50%
2
29-00
50%
D
(K6) EXTENDS THE RAM-AIR-TURBINE (RAT) WHEN AIRBORNE ABOVE 80 KNOTS WITH BOTH ENGINES N2 < 50% (K8)
2
30-32
50%
D
INHIBITS ANGLE-OF-ATTACK PROBE HEAT ON THE GROUND WITH EITHER ENGINE N2 < 50% (K7)
2
30-33
50%
D
INHIBITS THE AMBER LIGHT FOR THE TOTAL AIR TEMP PROBE ON THE GROUND WITH EITHER ENGINE N2 > 50%
2
80-11
52%
E
(K7) ILLUMINATES THE RESPECTIVE ENGINE START VALVE AMBER LIGHT WHEN THE STARTER CONTINUES TO OPERATE WITH N2 > 52% (K5)
STBY BUS ON
A
N2 SENSING
1
10 SEC
NVM CHANNEL 1
NORMAL MOMENTARY
CHANNEL 2 B B
TEST CH 2 NORMAL
L(R) ENG N2 SPEED CARD "SM"
POWER SUPPLY STATUS AND MAINT MSG PAGES
NON-MOMENTARY TEST NORM
TEST CH 1
N2 SENSING
CARD FRONT EDGE GND = NORMAL OPEN = FAULT
52/49% COMPARATOR
50/47% COMPARATOR
STARTER CUTOUT MESSAGE ECS HI FLOW INHIBIT EQUIPMENT COOLING LOAD SHED (GND SVC BUS) RAT EXTENSION AOA PROBE HEAT
ENGINE N2 SPEED CARD (P50)
ENGINE N2 SPEED CARDS B767-3S2F Page - 113
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PURPOSE
1
EICAS
N2 SPEED SENSOR
1
CHAN CHAPTERNO. SUBJECT N2
B767-3S2F Page - 114
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ENGINE INDICATION SYSTEM - CONDITION MONITORING General The condition monitoring system includes three pressure probes and one temperature sensor which send analog signals to the EEC. The EEC converts the converts the analog signal to digital data and sends a multiplexed signal to the PIMU. The ARINC communications and reporting system (ACARS) uses this information for diagnosis and fault information. The condition monitoring system includes signals from the following engine mounted sensors: • • • •
PS14 Fan Discharge Pressure P4.9 LPT Inlet Pressure T5 LPT Discharge Temperature P2.5 Compressor Inlet Pressure
ACARS
PS14 PS14 PROBE P25 PIMU P4.9 T5
P25 PROBE (P.A.RT OF T25/P2.5 SENSOR)
EEC
DFDAU
DFDR
P4.9 PROBE
T5 TEMPERATURE SENSOR
CONDITION MONITORING B767-3S2F Page - 115
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B767-3S2F Page - 116
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ENGINE INDICATION SYSTEM - PROPULSION INTERFACE MONITOR UNIT (PIMU) SYSTEM The Propulsion Interface Monitor Unit (PIMU) collects and stores fault information from the EEC. There are two PIMUs, 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 115vac ground service bus supplies power to the unit. Engine operating data is sent by both EEC channels. The unit accepts fault data from the EEC for 5 seconds after the airplane has landed and the air/ground relay has switched to the ground position. The monitor unit has a nonvolatile memory to store the data. The EICAS maintenance message "L(R) PIMU" appears if a fault is stored. The interface between the EEC and the aircraft components operate automatically. When the PIMU is interrogated, fault messages are shown on the face of the monitor unit. The PIMU interface buffer sends the data to the digital flight data acquisition unit (DFDAU) and the thrust management computer (TMC).
L ENG PIMU
NAMEPLATE
R ENG PIMU
MAIN EQUIPMENT CENTER ACCESS
24 CHARACTER LED ALPHANUMERIC DISPLAY
E2
E1
CHANNEL A CH A
INTERFACE BUFFER
EEC CH B
POWER SUPPLY
115V AC GND SVC
CHANNEL B
TMC
CHANNEL IN COMMAND
BIT
MONITOR VERIFY
DFDAU
CH A
RESET CH B
GND TEST
CH A AIR CH B
TEST LOGIC
ECS/MSG EICAS
L(R) PIMU
GROUND PIMU
PROPULSION INTERFACE MONITOR UNIT SYSTEM (PIMU) B767-3S2F Page - 117
MAINT RECALL
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PIMU
B767-3S2F Page - 118
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ENGINE INDICATION SYSTEM - 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 B data busses from the EEC. The flight is not finished at the time of landing. Thrust reverse, taxi and engine shutdown operations are yet to happen. The EEC will continue to monitor the system for faults. Any faults will be held in the EEC 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 procedures. 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.
IN CASE OF A REJECTED TAKEOFF, THERE IS NO AIR-TO-GROUND LANDING SIGNAL, SO THERE IS NO AUTOMATIC STORING OF EEC FAULTS BY THE PIMU.
FAULTS DETECTED BY THE EEC AFTER TOUCHDOWN WILL NOT BE STORED IN THE PIMU NVM.
EEC GETS POWER
AIRPLANE LANDS. FOR 5 SECONDS, THE PIMU STORES EEC FAULTS IN PIMU NVM.
FAULT MONITORING
FAULTS DETECTED BY THE EEC WILL BE AUTOMATICALLY RECORDED BY THE PIMU DURING THE FIRST 5 SECONDS AFTER TOUCHDOWN.
PIMU AUTOMATIC FAULT RECORDING DURING FLIGHT OPERATIONS B767-3S2F Page - 119
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ENGINE SHUTDOWN N2 3 SEC)
P34
SELF-TEST COMPLETE
NAMEPLATE NORM
END
EEC MAINT TEST (P61)
352 14-A
EEC CH A
EEC CH B
T-12
TEST IN
DATA
SENSOR
PROGRESS
BUS INOP
BITE
CHANNEL A CHANNEL B
115V AC GND SVCE
INSTRUCTION
BIT
BIT DATA RECEIVE MODE
MAINT RECALL
EEC
TMC DFDAU
P33
MONITOR VERIFY
AIR CH A
RESET CH B ECS/MSG
GND TEST
GND
EICAS
PIMU
PIMU BITE - MOST RECENT FLIGHT B767-3S2F Page - 121
GND TEST MODE
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L(R) PIMU DAVIN IS THE MAN
GND TEST FAIL
B767-3S2F Page - 122
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ENGINE INDICATION SYSTEM - PIMU GROUND TEST 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 the test. 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 115vac ground service bus must be powered. Operation Push the RESET switch to erase any faults stored in the PIMU non volatile memory. 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. If there are active faults detected by the EEC, they will be received by the PIMU and stored in non volatile memory. To view any faults that the PIMU has recorded in NVM, push the BIT switch once for each fault. If there are no faults or if you have viewed all the faults detected, the message END appears. To remove fault data from the PIMU, push RESET. This will erase PIMU NVM faults but will not erase the faults that are stored in the EEC.
115V AC GND SVCE P33 28V DC BAT BUS P34 APU/EXT PWR PNL
EEC MAINT L ENG POWER
NAMEPLATE
TEST NORM
TEST
1 EEC MAINT TEST (P61)
CH A
352 21-A
TEST IN
N1
PROGRESS
SENSOR
MOVE GND TEST TO CH A
2
PUSH BIT
CHANNEL A GROUND TEST (EEC POWERED)
BITE
INSTRUCTION
CHAN B DATA BUS
BIT
CH A PWR
MAINT RECALL
CH A MONITOR VERIFY
CH B
CH B PWR
CH A
1 RESET CH B
GND TEST
L ENG EEC EEC ALTERNATOR PIMU
PIMU BITE - GROUND TEST B767-3S2F Page - 123
INOP
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MOVE GND TEST TO CH B
CHANNEL B GROUND TEST (EEC NOT POWERED)
B767-3S2F Page - 124
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ENGINE INDICATION SYSTEM - 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 flight legs. The maintenance recall procedure will transfer faults only for the channel in control of the engine at that time. The engine must be shut down 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, un-power that EEC by cycling the maintenance ground test switch to NORM, then back to the TEST position, and finally pull the appropriate engine channel circuit breaker. This procedure changes the channel-in-control as shown on the EPCS EICAS page. 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. 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.
TEST
TEST
IN
READY
IN
PROGRESS 1
PUSH & HOLD: MONITOR VERIFY
RELEASE: MONITOR VERIFY
3
4
PUSH: MAINT RECALL
FLIGHT
350 $27-A
TEST
FLIGHT
LEG
NO 28V DC
IN
LEG
DETECTED
PROGRESS
1
5
9
2
PROGRESS
6
PUSH: BIT
7
PUSH: MAINT RECALL
1
8
351 $26-A
TEST
FLIGHT
EXITING
R ADC
IN
LEG
MAINT
CHANFAIL
PROGRESS
PUSH: BIT
10
MODE
PUSH: MAINT RECALL
11
PIMU MAINTENANCE RECALL B767-3S2F Page - 125
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12
PUSH: MONITOR VERIFY
B767-3S2F Page - 126
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ENGINE INDICATION SYSTEM - ELECTRONIC PROPULSION CONTROL SYSTEM (EPCS) General The values for various engine controls and status parameters appear on the EPCS maintenance pages 1 and 2. The parameters are shown as real time. AUTO EVENT or MAN EVENT data. EPCS Page 1 Data from both channels of the EEC on each engine appear. The channel which is currently in control of the engine operations is indicated by a square around the channel letter. In the case of the AUTO / MAN EVENT the square displayed indicates the channel which controlled that engine at the time the event was recorded. EPCS Page 2 Page 2 of the EPCS display is accessed by pressing the EPCS maintenance switch a second time. Page 2 is real time information only. There is no MAN / AUTO EVENTS for this page. The hexi-decimal ARINC 429 labels can be decoded using the FIM manual, with the PIMU MESSAGE INDEX.
EPCS ____
EPCS ____ PAGE 1 A
1.6 99.8 34.5 4 14.5 0 0 0.0 0.0 7 20 14 OIL T YEL
B
1.6 99.7 34.4 4 14.5 1 0 0.0 0.0 7 20 15
A VSV VBV TRA T 12 P0 HPTC LPTC T/R L T/R R T25 T3 PS 3
1.4 100.0 34.0 4 14.5 0 0 0.0 0.0 7 20 14
AUTO EVENT
PAGE 2
B
A
1.4 100.0 34.1 4 14.5 1 0 0.0 0.0 7 20 15
0840 0300 0802 4000 0E01 4140 1180
B
0800 0300 6802 4000 0E01 4140 1180
LABEL 270 271 272 273 274 275 276
EGT RED
EICAS ELECTRONIC PROPULSION CONTROL SYSTEM PAGE B767-3S2F Page - 127
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A
0801 0300 6802 4000 0E01 4240 1180
B
0841 0300 0802 4000 0E01 4240 1180
B767-3S2F Page - 128
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ENGINE CONTROL - CLUTCH AND MICROSWITCH PACKS General The autothrottle clutch pack assembly is the interface between the autothrottle system and the engine fuel control system. It is 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 below the drum. Autothrottle Clutch Packs The autothrottle clutch packs supply friction and feel for the thrust levers (manual) and let the autothrottle servo unit move the thrust levers. The clutch packs are on a common shaft. The thrust levers connect to one face of a clutch pack. The autothrottle servo unit connects 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. In reverse thrust, the autothrotle clutch cannot increase engine thrust. In reverse thrust, all thrust changes are manual. The clutch packs make manual override of the servo unit possible at all times. 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 operates the switches to send thrust lever position signals to other aircraft systems. Training Information Point The 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 adjusted to have all switches in the group operate at the same time. In addition, there is an adjustment bolt for each group. Adjust the bolt to get the switches to operate at the correct thrust lever angle.
To adjust the switch group, put the thrust levers at the proper angle as described in the Maintenance Manual. A scale on the forward drum shows the position. Push on the lock channel to disengage the adjustment bolt. Turn the bolt to adjust the switch. Make sure the position is correct by a continuity test on the applicable pins in the electrical connector. When the position is correct, release the lock channel to re-engage the bolt. Switches These are the switches: • • • • • •
S1, S5 - L/R LANDING WARNING S2, S3 - L AUTOBRAKE/AUTOBRAKE REJECTED TAKEOFF (RTO) S6, S7 - R AUTOBRAKE/AUTOBRAKE REJECTED TAKEOFF (RTO) S8, S11 - L/R THRUST REVERSER DIRECTIONAL CONTROL VALVE S10, S14 - L/R SPEEDBRAKE RETRACT S12, S16 - L/R THRUST MANAGEMENT SYSTEM (TMS) THRUST REVERSE • S17 - LOAD SHED/PRESSURE CONTROL L • S18 - LOAD SHED/PRESSURE CONTROL R.
CONTROL RODS TO THRUST LEVERS AUTOTHROTTLE CLUTCH PACK S14
S5
CLUTCH LINK AUTOTHROTTLE SERVO UNIT
S18 S16
S6 SPACER S7
S8 S10
S2 S1 SPACER S3
S17 S12 S11 FWD
CAM FOLLOWING ARMS
FWD
MICROSWITCH PACK
MICROSWITCH ASSEMBLY
FWD DRUM
MOUNTING SCREWS MOUNTING ARM
SWITCH CAM CAM FOLLOWING ARMS
SHAFT
SWITCH
MICROSWITCH PACK
ENGINE CONTROL - CLUTCH AND MICROSWITCH PACKS B767-3S2F Page - 129
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LOCK CHANNEL ADJUSTING BOLT
B767-3S2F Page - 130
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ENGINE CONTROL - THRUST LEVER ANGLE (TLA) RESOLVERS General 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 on 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. Each EEC channel sends a sine wave signal through its respective connector to the rotor of the dual coil TLA resolver. The excitation induces a sine-cosine feedback signal for each channel as the rotor moves in response to power lever position changes. The EEC converts the sensed analog feedback signals into a digital thrust lever angle value. The EEC uses this phase angle to determine commanded N1.
AUTOTHROTTLE CLUTCH PACK ASSEMBLY
TLA RESOLVER FORWARD ACCESS DOOR
AUTOTHROTTLE SERVO MOTOR
THRUST LEVER CONTROL RODS CLUTCHES
RESPONSE SIGNALS SENSING CIRCUITS POWER SUPPLY CHANNEL A CHANNEL B
STATORS ROTORS EXCITATION SIGNALS POWER SUPPLY
SENSING CIRCUITS TLA RESOLVER
RESPONSE EEC SIGNALS
TLA RESOLVER LINKS (2)
CHANNEL A CONNECTOR EEC EXCITATION
PHASE ANGLE
THRUST LEVER ANGLE RESOLVER (2) AUTOTHROTTLE SERVOMOTOR CHANNEL B CONNECTOR
ENGINE CONTROL - THRUST LEVER ANGLE (TLA) RESOLVERS B767-3S2F Page - 131
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ENGINE CONTROL SYSTEM - FADEC SYSTEM DESCRIPTION
• Auto-Throttle System (ATS)
General
It is extensive information processing capabilities, more than any other, that distinguishes FADEC from mechanical engine control systems.
The General Electric CF6-80C2 full authority digital electronic control (FADEC) system 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 the electronic engine control (EEC). The FADEC system is divided into subsystems to perform two basic functions - information processing and engine control.
ENGINE CONTROL refers the FADEC's ability to physically control the operating, performance and efficiency characteristics of the engine. Capabilities in this area include precise control over fuel flow, primary and parasitic airflow, internal rotor to stator clearances (Active Clearance Control), engine start sequencing and igniter operation.
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. This information comes form engine control switches in the flight deck, thrust lever position inputs, temperature and presser inputs on the engine. The EEC uses this information to control the engine through the EEC which also sends data and messages to EICAS, the SEI and the engine discrete card. The flight management computer (FMC), thrust management computers (TMC) and the air data computers (ADC) also interface with the EEC. The engine control functions control the engine fuel and air systems to operate the engine efficiently at all rated performance levels. The FADEC system is composed of an engine control (EEC), Hydro-Mechanical Unit (HMU), Permanent Magnet Alternator (PMA), Engine rating Plug, Engine Identification Plug, engines sensors and components from the Variable Stator Vane (VSV), Variable Bleed Valve (VBV), HPT Active Clearance Control (HPTACC) and Engine Starting and Ignition systems. It is divided into seven separate subsystems that provide two basic system functions - Information Processing and Engine Control: • Information processing refers to the FADEC's ability to input, manipulate and output large amounts of electronic data. Using these functions, the FADEC computer gathers information about the environment and operating conditions within the engine. With the information, the computer calculates fuel and air flows required to maintain engine operation at the rated performance levels with peak efficiency. Information processing also allows the FADEC computer to communicate directly with other computerized aircraft systems including the: • Engine Indicating and Crew Alerting System (EICAS) • Air Data Computer (ADC)
EICAS (CHAPTER 31)
SEI (CHAPTER 77)
FMC (CHAPTER 34)
TMC (CHAPTER 22)
ADC (CHAPTER 34)
. EEC DISCRETES CARD (CHAPTER 73)
IDLE SIGNAL
PNEUMATIC DEMAND
THRUST LEVERS (CHAPTER 76) FUEL CONTROL SWITCHES (CHAPTER 73) TLA RESOLVER (CHAPTER 73)
MICROSWITCH PACK (CHAPTER 22)
EEC (CHAPTER 73)
CONTROL ALTERNATOR (CHAPTER 73)
THRUST REVERSER (CHAPTER 78)
T12 SENSOR (CHAPTER 73)
P25/T25 SENSOR (CHAPTER 73)
FADEC SYSTEM DESCRIPTION B767-3S2F Page - 133
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HYDROMECHANICAL UNIT (HMU) (CHAPTER 73)
.
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ENGINE CONTROL SYSTEM - ELECTRONIC ENGINE CONTROL (EEC) The electronic engine control (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 robes 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 Identification Plug on connector J15 programs the EEC for 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 on the P61 panel is in the TEST position. 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), onengine components (J7-J13) and EEC use (J14-J15).
Aircrft Interface Connectors (J1-J6) • J1 Ignition Exciter #1. DC Power In/Out; Channel A Ground Handeling Bus Power In • J2 Ignition Exciter #2. DC Power In/Out; Channel B Ground Handeling Bus Power In • J3 Fuel On; Starter Air Valve Open; Chanel A Reset: EEC Fault; Digital Data Bus (ADC & TMC) In/Out, Channel A TLA resolver In/Out • J4 Single/Dial; Igniters; Idle Select; Hard Reversionary Mode; Channel B TLA Resolver In/Out • J5 Aircraft Type; Engine Position (L/R); channel A Thrust reverser Position • TMC Disconnect; Operating Mode Select (Control or Test); Channel B Thrust Reverser Position
Engine Interface Connectors • • • • • • • • • • •
J7 Black Channel A J8 Brown Channel B N2 Sensor; ESCV Solenoid, Escv Position Switches; HMU J9 red Channel A J10 Orange Channel B Control Alternator; Starter Air Valve; N1 Sensor; T12 J11 Yellow Channel A J12 Green Channel B T2.5; HPTC Valve; VSV Actuators; VBV Actuators J13 Blue Channel A and B T3; T49; T5; Engine Oil Temperature Sensort; Fuel Flow Transmitter
ENGINE RATING PLUG CONNECTOR (J15)
EEC
ENGINE DENTIFICATION PLUG
FWD PS3 ENGINE RATING PLUG
ENGINE CONTROL SYSTEM - ELECTRONIC ENGINE CONTROL (EEC) B767-3S2F Page - 135
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SERIAL NUMBER PLUG CONNECTOR (J14)
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ENGINE CONTROL SYSTEM - ELECTRONIC ENGINE CONTROL (EEC) (CONT) Data Plugs • J15 Engine Rating Plug • J14 Identification Plug These two plugs are captive to the engine by lanyards. Multiple tables are contained in the EEC and the P14 determines the rating table to be used. The P15 provides engine hardware informatin to the EEC: • • • •
N1 Modifier EGT Shunt Valve Active Clearance Control Schedules Engine Serial Number (Programed Through J15)
Pressure Inputs The EEC has pressure transducer and signal conditioning circuits. The pressures measured are as follows: • 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.
ELECTRONIC ENGINE CONTROL SWITCHES (P5)
ENGINE RATING PLUG CONNECTOR (J15)
EEC
ENGINE IDENTIFICATION PLUG
FWD PS3 ENGINE RATING PLUG
ELECTRONIC ENGINE CONTROL (EEC) (CONT) B767-3S2F Page - 137
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SERIAL NUMBER PLUG CONNECTOR (J14)
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ENGINE CONTROL SYSTEM - CONTROL ALTERNATOR Purpose The control alternator provides the EEC channels A and B with electrical power. Characteristics The alternator is located on the forward center section of the accessory gearbox. Opening the thrust reverser allows access. The alternator consists of two separate assemblies: • Rotor • Stator Rotor The rotor is a permanent magnet assembly - Permanent Magnet Alternator PMA. It is mounted to the Accessory Gearbox (AGB) splined drive shaft with a lock nut. Stator The stator mounts on the AGB case with three bolts. The stator has three independent windings. Two windings power the EEC channels A and B. Operation The alternator operates whenever the gearbox is turning. It will meet all required EEC power at 11% N2. It continues to meet the power requirements until the N2 decreases below 9%. If one phase of either or both windings fail, the control alternator continues to meet all EEC power requirements if the N2 is above 45%.
PERMANENT MAGNETS
WINDINGS (2) ROTOR NUT
FLATS (3) AGB DRIVE SHAFT MOUNTING PAD CHANNEL A O-RING
CHANNEL B STATOR
PERMENANT MAGNET ALTERNATOR B767-3S2F Page - 139
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ENGINE CONTROL SYSTEM - INLET SENSORS (T12) Engine Inlet Temperature Sensor (T12) There are two T12 Inlet temp sensors. Each supplies inlet temp 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 elements in the sensor are resistive thermal devices. Hence, temperature changes in the engine inlet area varies the resistance of the probes. The housing the sensor is mounted in protects it from physical damage. It also prevents water and ice contact interfering with the accurate operation of the probe. The T12 sensor is used by the EEC to correct N1 and N2 speed inputs, and to calculate the position of the Fuel Metering Valve and the HPTACC Valve. Inputs from the sensor mounted in the 2:00 position are received and processed by Channel A, and channel B inputs are from the sensor mounted at the 10:00 position. Each EEC channel supplies a 10 ma direct current excitation signal to its respective sensor. The voltage drop across the sensor is measured by the EEC and corrected for ram air effects to determine the inlet air temperature. The digital equivalent of each input is made available at the aircraft interface for monitoring.
I +V -
T
ELECTRICAL CONNECTOR
CHANNEL A
10:00
CHANNEL B
I PROTECTIVE HOUSING
PLATINUM WIRE ELEMENT
AIRFLOW
T
2:00
ELECTRICAL TEMPERATURE SENSORS (T12) B767-3S2F Page - 141
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+ V -
EEC
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ENGINE CONTROL SYSTEM - INLET SENSORS (P/T 2.5) General The P2.5 probe is a part of the compressor inlet temperature/pressure T2.5/ P2.5 sensor. The P2.5 probe senses the total pressure of the high pressure compressor inlet airflow. The T2.5/P2.5 sensor is on the fan frame hub outer surface at the 7:30 position. The P2.5 probe has a pitot tube to sense pressure. The pressure signal goes to a P2.5 pressure transducer in the EEC. The operation range of the P2.5 input to the EEC is from 2 to 75 psia.
Compressor Inlet Temperature/Pressure Sensor (T2.5) The compressor inlet temperature sensor (T2.5), is part of the T2.5/P2.5 temperature sensor. This sensor is mounted on the fan frame at the 7:30 position between the number 8 and 9 fan struts. The sensor has two separate temperature sensing elements, one for each channel of the EEC. Once again temperature varies resistance in this sensor and that change is read by the EEC as a temperature. The T2.5 is used by the EEC to correct N2 speed inputs. Two T2.5 inputs are received from the sensor. One input is received and processed by Channel A, and the other by Channel B. Each channel supplies 10 ma (max) direct current excitation signal to the sensor. The digital equivalent of each input is made available at the aircraft interface for monitoring. Note:
The P2.5 portion of this sensor is not currently used.
P2.5 PORT
FAN STRUT 8 T2.5 CONNECTORS
FWD
ELECTRICAL TEMPERATURE SENSORS (P/T 2.5) B767-3S2F Page - 143
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ENGINE CONTROL SYSTEMS - EEC DISCRETES PRINTED CIRCUIT CARD One EEC discrete’s 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 in puts and outputs. The card has two sections, one for each engine. The 28vdc battery bus and the left 28vdc bus supply power to the card's left engine section. The 28vdc battery bus and the right 28vdc bus supplies power to the card's right engine section. CAUTION: THIS CARD IS STATIC SENSITIVE. DO NOT HANDLE BEFORE READING THE PROCEDURE FOR HANDLING ELECTROSTATIC DISHARGE SENSITIVE DEVICES (REF 2041-01). THE CARD CONTAINS DEVICES THAT CAN BE DAMAGED BY STATIC DISCHARGE.
Characteristics The card is a printed circuit type. Relays on the card provide interface between inputs and outputs. The card has two sections, one for the left engine and one for the right. The left engine section is shown.
Power The left engine section of the card is powered by the 28 volt dc battery bus and the left 28 volt dc bus, the right engine section is powered by the 28 volt dc bat. bus and the right 28 volt dc bus.
28V DC R BUS
RIGHT ENGINE SECTION (SAME TO LEFT)
28V DC BAT BUS
EEC DISCRETES PRINTED CIRCUIT CARD
28V DC L BUS
P50 8
6
7
10
6
5
4
3
2
1
6
5
4
3
2
1
POWER L ENG EEC DISCRETES
P11 5 SEC
T / D
+ -
TO IDLE SELECT CONTROL CIRCUIT 10
3
8
TIME DELAY (K12) L ENG ANTI-ICE
K1 COWL ANTI-ICE
AIR HYD PUMP OVERSPEED CONT CARD
K4 ADP
BLEED STATES
AIR SUPPLY ISLN VLV CLOSED IND
K10 R ISLN VLV
L PACK FLOW
K3 L ECS HI/LO
CONT CARD
TO FMC
TO FMC
EEC
K2 L ECS ON/OFF
P5 SWITCHES
TMC
P50 CARD FILE (MEC)
28V DC BAT BUS APU ENG START/ECS DISCRETE P11
INBD EEC DISCRETES PRINTED CIRCUIT CARD (P50)
BLEED AIR CTR ISN VALVE
AUX POWER CONTROL UNIT
EEC DISCRETES PRINTED CARD B767-3S2F Page - 145
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ENGINE CONTROL SYSTEM - HMU FUEL METERING OPERATION General Fuel flow is metered by the hydro-mechanical 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.
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 Fuel Metering Valve A fuel metering valve (FMV) inside the HMU controls fuel flow to the fuel nozzles. The hydraulically driven metering valve is controlled by the FMV EHSV. Control of the EHSV is through two coils , one for each EEC channel. The controlling EEC channel increases current through its EHSV coil to hydraulically open the FMV. The FMV has two position indicating resolvers, each providing feedback to and getting power from it’s own respective EEC channel.
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 volt battery bus. It has two latching coils: • Run • 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.
The bypass valve has a piston inside a multi ported sleeve. Un-metered 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. The delta P regulator maintains a constant pressure drop across the FMV. This makes the fuel flow rate vary with the FMV position. The regulator monitors the pressure difference between the un-metered fuel input and the 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.. .
VBV SERVO FUEL PORTS
HPTC REFERENCE PRESSURE PORT ACCESSORY GEARBOX FORWARD SIDE
HPSOV SOLENOID CONNECTOR
SERVO FUEL INLET
EHSV (5)
HPSOV POSITION SWITCH CONNECTOR EEC CHANNEL A CONNECTOR
EEC CHANNEL B CONNECTOR
VSV SERVO FUEL PORTS
FUEL DISCHARGE (HIDDEN)
FUEL INLET
HPTC VALVE PORT
TOP
FWD
TOP DRIVE COUPLING
BYPASS FUEL RETURN RIGHT, TOP SIDE
LEFT, BOTTOM SIDE
ENGINE CONTROL SYSTEM - HMU FUEL METERING OPERATION B767-3S2F Page - 147
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ENGINE CONTROL SYSTEMS - HYDROMECHANICAL UNIT (HMU) (CONT) The fuel metering system 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 electrical connectors for electrical interfaces with the aircraft and MU with the fuel pump and nozzles. There are five hydraulic connections for control interface 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 are: • • • •
Fuel inlet from the fuel pump Fuel discharge to the fuel nozzles Fuel bypass discharge to the fuel pum Servo fuel inlet from the servo fuel heater
The hydraulic connections are: • • • • •
Servo fuel pressure to the Low Pressure Turbine Cooling Valve (LPTC) Servo fuel pressure to the High Pressure Turbine Cooling Valve (HPTC) 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)
Note:
The LPTC system is currently not used on the 767. The EHSV is still located on the HMU, however the control valve has been removed. The system flows constantly without and external control systems.
The electrical connections to the HMU are: • • • •
Fuel control signals from the EEC channel A Fuel control signals from the EEC channel B HPSOV solenoid inputs from the fuel control valves HPSOV position indicating outputs to the EEC
VBV SERVO FUEL PORTS
HPTC REFERENCE PRESSURE PORT ACCESSORY GEARBOX FORWARD SIDE
HPSOV SOLENOID CONNECTOR
SERVO FUEL INLET
EHSV (5)
HPSOV POSITION SWITCH CONNECTOR EEC CHANNEL A CONNECTOR
EEC CHANNEL B CONNECTOR
VSV SERVO FUEL PORTS
FUEL DISCHARGE (HIDDEN)
FUEL INLET
HPTC VALVE PORT
TOP
FWD
TOP DRIVE COUPLING
BYPASS FUEL RETURN RIGHT, TOP SIDE
LEFT, BOTTOM SIDE
ENGINE CONTROL SYSTEMS - HYDROMECHANICAL UNIT (HMU) (CONT) B767-3S2F Page - 149
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ENGINE CONTROL SYSTEM - HMU FUEL METERING OPERATION (CONT) Overspeed Governor The overspeed governor senses N2 speed through the HMU mechanical drive from the accessory gearbox. If the N2 exceeds 113.4 percent, the governor overides the delta P regulator input to the bypass valve to reduce the 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 S/O 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% 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.
CONTROL INPUT FROM EEC
FEEDBACK TO EEC FIRE CUTOFF
UNMETERED FUEL METERED FUEL SERVO FUEL
METERING VALVE EHSV
IN FROM FUEL PUMP
METERING VLV RESOLVERS
NORM
RUN FUEL CONT SW (P10)
FIRE SW (P8)
HPSOV
OUT TO NOZZLES
PRESSURIZING AND SHUTOFF VALVE
METERING VALVE
28V DC BAT BUS
HPSOV SOLENOID
VALVE POSITION SWITCH
ENG VALVE P10
DIFFERENTIAL PRESSURE REG/ BYPASS VALVE
RETURN TO FUEL PUMP INTERSTAGE
L (R) ENG FUEL VAL (C)
ACCESSORY GEARBOX
A B
N2 OVERSPEED GOVERNOR
PRIMARY ENGINE DISPLAY
L (R) ENG O/S GOV (S,M)
O/S SWITCH
STATUS OR ECS/MSG PAGE
HYDROMECHANICAL UNIT (HMU)
EEC
HMU FUEL METERING OPERATION B767-3S2F Page - 151
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EICAS
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ENGINE CONTROL SYSTEM - 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 the Permanent Magnet Alternator (PMA) when the engine is running, and from the aircraft when the engine is not running. The EEC sends analog output signals to the hydro-mechanical 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 T12 or PO sensor data is invalid. If crosschannel data is invalid, the EEC switches to the soft reversionary mode. • Comparisons are made between N1, N2, P3 or T2.5 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, CCC valves and the thrust reverser interlocks fail-safe to open or closed. • The EEC uses 28vdc aircraft power if power is not available from the control alternator.
COMMAND P0
ENGINE AIR SYSTEMS
FEEDBACK
PNEUMATIC PS3 COMMAND T12 SENSOR
T25 T3 OIL TEMP (TEO) T49 (EGT) N1 N2
P25/T25 SENSOR ENGINE ANALOG
HMU
ENGINE AIR SYSTEMS FEEDBACK AIRCRAFT DIGITAL AIRCRAFT ANALOG
METERED FUEL FLOW
FEEDBACK
T12
ENGINE AIR SYSTEMS
ADCS (ALT, TAT, CAS, PT, T STATIC)
STANDBY ENGINE INDICATOR
TMC (BLEED DEMAND, N1 TRIM) TLA RESOLVER T/R POSITION
THRUST LEVER
T/R INTERLOCK
AIRCRAFT ID/ENG LOCATION EICAS (N1, N2, EGT, EEC STATUS, & FAULTS)
ENGINE RATING PLUG
AIRCRAFT DISCRETE
EEC DISCRETES
APPROACH IDLE
TEST SW
TEST
FUEL CONTROL SW
RESET
EEC CONTROL SW
HARD REV MODE
ALTN MODE INDICATION
START
START/IGN SW POWER
PIMU
CONTROL ALTERNATOR
STARTER AIR VALVE
CROSSTALK
AIRCRAFT POWER
CHANNEL A
POWER
IGNITORS
CHANNEL B CHANNEL B INPUTS SAME AS CHANNEL A
CHANNEL B OUTPUTS SAME AS CHANNEL A EEC
EEC INPUTS / OUTPUTS B767-3S2F Page - 153
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MULTIPLE ANALOG SIGNALS
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ENGINE CONTROL SYSTEM - 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 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 percent). Aircraft power is supplied when: • the engine is being started • the engine fuel control switch is in the RUN position • the EEC maintenance engine power switch 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, the channel in control of the engine 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.
A
MEMORY
28V DC L BUS
DATA INTERFACE
TO AIRPLANE
SOLENOID DRIVER
TO ENGINE (ACTIVE CHANNEL ONLY)
PWR SUPPLY
TEST K1169 L ENG PWR CH A
EEC MAIN TEST (P61)
START
RECTIFIER
FROM AIRPLANE SYSTEMS
DIGITAL INTERFACE
FROM ENGINE SENSORS
SIGNAL COND
CPU
K11736 ENG
START 3 RLY RUN
CHANNEL A K1036 CH A RST RLY (P36)
CHANNEL B
SIGNAL COND
PRESS XDCR
SIGNAL COND
PRESS XDCR
PRESSURE SENSORS
CROSSTALK
A CONTROL ALTERNATOR
EEC
EEC OPERATION B767-3S2F Page - 155
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PRESSURE SIGNAL INPUTS
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ENGINE CONTROL SYSTEM - 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. If a channel reset signal is received 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 the active channel.
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. The EEC will then send signals to the solenoid valve inside the HMU to latch open the Pressurizing and Shutoff Valve. When the fuel control switch is set to the CUT-OFF position a signal is sent to the EEC and it signals the latch closed solenoid in the HMU to close the Pressurizing and Shutoff Valve. The fire switch pulled up to the FIRE position will also signal the EEC to close the Pressurizing and Shutoff Valve.
28V DC L BUS L ENG EEC PWR CH A L ENG EEC PWR CH B
28V DC
1
A K1169 L ENG PWR CH A (P36)
P11
POWER FUEL ON
FIRE
28V DC BAT
RESET
RUN
L ENG FUEL CONTROL VALVE RESET A
K1036 L ENG CH A RST (P36)
P11
TO FUEL/IGNITION CONTROL RELAY(S)
CHANNEL A CHANNEL B
28V DC HOT BAT L SPAR VALVE RESET B
NORM FIRE SWITCH (P8)
P6
CUTOFF FUEL ON
FUEL CONTROL SWITCH (P10)
COMMON RETURN RESET
1 1
28V DC - ENG STARTING (N2 50%
ON
1
N1 CMD > 1.02 (N1 MAX)
S1 L ENG EEC CONTROL SW (P5)
L ENG N2 SPEED CARD (P50)
POWER AND MODE SELECT
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L ENG EEC
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ENGINE CONTROL SYSTEM - 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 LADC, T2 RADC, T12 CH A, and T12 CH B) to select a T2 value for calculating N1 command. The ambient pressure inputs (PO LADC, PO RADC, 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 LADC and PT2 RADC). 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.
Soft Reversionary Control Mode The normal control mode is used if PT2 LADC and PT2 RADC 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. 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, the ALTN lights on the EEC switches comes 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 corner point DTAMB value. The thrust can increase by using the corner point DTAMB value instead of the DTAMB value used in the soft reversionary control mode. This can cause over boosting of the engine depending on actual ambient conditions and thrust lever angle. To prevent over-boosting, the thrust levers must be pulled back to an intermediate position prior to selecting the hard reversionary control mode. The corner point DTAMB value is used to calculate an assumed TAMB as altitude changes, and to calculate Mn and Q. N1 command is calculated using the calculated values for Mn, Q, TAMB and DTAMB and the PO, T2, TLA and bleed values.
PT2 (L ADC) PT2 (R ADC)
TO EEC SW ALTN LIGHT AND EICAS
FAULT LOGIC
PO (CH A), PO (CH B) PO (L ADC), PO (R ADC) T12 (CH A), T12 (CH B) T2 (L ADC), T2 (R ADC) BLEED VALVES (TMC) TLA
NORMAL CONTROL
ALTERNATE MODE SELECT (USING EEC SWITCH)
PT2 INPUT FAIL LAST VALID ADC DATA
THRUST RATING LOGIC
SOFT REVERSIONARY CONTROL
N1 CMD TAT/T12
CORNERPOINT AT AMB (30C)
HARD REVERSIONARY CONTROL
PO PS3 N2 MIN MIN IDLE SEL
CPU
FROM L ADC
FROM R ADC
PO
CROSS CHANNEL DATA BUS
DIGITAL INTERFACE
T2
PT2
DIGITAL INTERFACE
T2 PT2
N1 CPU
N2
T/R POS
ACCEL/ DECEL SCHEDULE
EEC
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FUEL FLOW
LIMIT PROTECTION TLA
CHANNEL A CHANNEL B
PO
N1 N2 PS3
N2 IDLE
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TR MAX
REVERSE CONTROL
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ENGINE CONTROL SYSTEM - CONTROL MODES (CONT) 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 430psid. The N2 limit schedule is used in addition to a mechanical overspeed governor in the hydro-mechanical 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 flame out. This control ensures that all engines respond to thrust lever angle changes at the same rate.
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 percent). 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.
Reverser 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.
PT2 (L ADC) PT2 (R ADC)
TO EEC SW ALTN LIGHT AND EICAS
FAULT LOGIC
PO (CH A), PO (CH B) PO (L ADC), PO (R ADC) T12 (CH A), T12 (CH B) T2 (L ADC), T2 (R ADC) BLEED VALVES (TMC) TLA
NORMAL CONTROL
ALTERNATE MODE SELECT (USING EEC SWITCH)
PT2 INPUT FAIL LAST VALID ADC DATA
THRUST RATING LOGIC
SOFT REVERSIONARY CONTROL
CORNERPOINT AT AMB (30C)
N1 CMD TAT/T12 PO PS3 N2 MIN MIN IDLE SEL
HARD REVERSIONARY CONTROL
CPU
FROM L ADC
FROM R ADC
PO
CROSS CHANNEL DATA BUS
DIGITAL INTERFACE
T2
PT2
DIGITAL INTERFACE
T2 PT2
N1 CPU
N2
T/R POS
ACCEL/ DECEL SCHEDULE
TR MAX
EEC
ENGINE CONTROL SYSTEM - CONTROL MODES (CONT) B767-3S2F Page - 163
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FUEL FLOW
LIMIT PROTECTION TLA
CHANNEL A CHANNEL B
PO
N1 N2 PS3
N2 IDLE
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REVERSE CONTROL
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ENGINE CONTROL SYSTEM - ENGINE IDLE SELECT 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 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 pressure regulating and shutoff valve (T/R 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. 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 degrees after having been below 23 degrees. • 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 two EEC’s do not agree, and 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.
FADEC engines are susceptible to flameout at minimum idle when encountering inclement weather. The ignition select switch is used to comand approach idle preventing possible flameout.
A 28V DC L BUS L ENG IDLE CONTROL P11
28V DC GND HANDLING BUS ENG IDLE CONTROL
4
2
1
K1034 L T/R VALVE RELAY (P36)
K434 L IDLE SOL RLY (P36)
K1025 L T/R DEPLOY IDLE RLY (P36)
TO RIGHT ENGINE CIRCUIT 28V DC R BUS
A
GND AIR
T/D 5 SEC
EICAS
MIN IDLE
P34
COMMON RETURN CHANNEL A CHANNEL B
AIR GND
R ENG IDLE CONTROL
FLAPS LANDING
P11
1
ENERGIZED WITH T/R PRSOV
2
ENERGIZED WHEN T/R DEPLOYED AND FIRE HANDLE NORMAL
3
IDLE DISAGREE (C) MESSAGE
4
ENG LOW IDLE (C) MSG - N1 BELOW APPROACH IDLE - TAI ON
K167 SYS 1 AIR/GND BAT RLY (P36)
EEC DISCRETES CARD (P50)
L FLAP/STAB POSITION MODULE (P50)
TAI ON
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K141 SYS AIR/ GND RLY (P36)
CONTINUOUS IGNITION ON K785 L ENG TAI IDLE (P36)
ENGINE IDLE SELECT Page - 165
3
R ENG
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EEC (L ENG)
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ENGINE START SYSTEM - START SYSTEM AIR SOURCES Ground Air Ground air is available through the ground service pneumatic connections. The nominal required pressure is 45 psi.
APU Air The auxiliary power unit (APU) provides approximately 54 psi air. The APU air supply shutoff valve (SOV) is controlled by APU switch on the P-5 overhead panel. The center isolation valve is normally open. The left and right isolation valves are controlled by switches on the P-5 overhead panel. During a main engine start the APU operates at a higher speed to insure adequate air flow.
Engine Air During a cross-engine start, air from an operating engine is used to start the other engine. Two engine air sources are available; 8th stage bleed air and 14th stage bleed air. At high engine speeds, the high pressure SOV is closed and 8th stage air is used. At low engine speeds (idle to 75% N2), the high pressure SOV is open, the low pressure air supply check valve is closed, and 14th stage air is used.
General Operation During a cross-engine start, the air supply pressure regulating and shutoff valve (PRSOV) must be open on the running engine and closed on the engine that is being started. The PRSOV is controlled by switches on the P5 overhead panel. To pressurize the starting system, the air conditioning pack control selector must be in "OFF", the pneumatic starter control valve must be open and applicable PRSOVs (depending upon the air source) are shut. The pneumatic starter control valve is controlled by the engine start switch on pilots' overhead panel.
LEFT ISOLATION VALVE
RIGHT ISOLATION VALVE
E GROUND AIR SOURCE
LEFT PRSOV
F
RIGHT PRSOV
D
R ENGINE
A
PRESSURE REG VALVE (PRV)
8TH STAGE SUPPLY CHECK VALVE
B
CENTER ISOLATION VALVE
AIR SUPPLY PRECOOLER
APU AIR SUPPLY VALVE
C
TO R ENGINE START CONTROL VALVE
APU
8
PNEUMATIC STARTER CONTROL VALVE
14
60 40
G
L ISLN
START CONTROL VALVE DISAGREEMENT LIGHT
80
L R
HIGH PRESSURE VALVE
DUCT PRESS PSI
20
0
VALVE
R ISLN VALVE
E
F C ISLN
ENG START L
SINGLE
V A L V E
DUCT LEAK
R
BOTH
VALVE
VALVE
BLEED
STARTER
DL UE CA TK
HI STAGE GND
IGNITION/START CONTROL SWITCH
AUTO OFF
GND
AUTO OFF
CONT
CONT
FLT
FLT
G PILOTS OVERHEAD PNL (P5)
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BLEED
ADP APU
L ENG O
AF F
ENGINE START SYSTEM AIR SOURCES
DUCT LEAK
D
V A L V E
C
OVERHEAD PANEL (P5)
HI STAGE
R ENG O
BFF
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ENGINE START SYSTEM - START SYSTEM COMPONENTS Location and Features Pneumatic Starter - The pneumatic starter is mounted to the accessory gearbox in the 6 o'clock position. It provides the initial rotation of the N2 compressor needed to ensure a successful engine start. Pneumatic Starter Control Valve - The pneumatic starter control valve is mounted between the starter inlet and the air supply ducts and controls the flow of air to the pneumatic starter. Engine Ignition and Start Control Module - The engine ignition and start control module located on the P5 overhead panel provides a means of controlling starting operations. The module contains two valve lights, the ignition selector switch and the two engine start switches. The operations of the switches pertaining to engine ignition are discussed in the Engine Ignition Chapter.
STARTER (REF)
1
1
MANUAL DRIVE ACCESS
PACKING FILTER ELEMENT
STARTER CONTROL VALVE
FILTER SPRING CAP THRUST REVERSER LATCH ACCESS DOOR
ENGINE START SOLENOID
FROM EEC VALVE BODY
ENGINE START SOLENOID
ACTUATOR POSITION INDICATING SWITCH ASSEMBLY
START SYSTEM COMPONENTS B767-3S2F Page - 169
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ENGINE START SYSTEM - ENGINE START CONTROL EICAS R (L) Engine Starter Message The EICAS level C message, L(R) ENG STARTER is displayed after a 5 second time delay if the starter valve does not open when commanded.
EICAS R (L) Starter Cutout Message If the starter valve fails to close, or if K666 does not relax before N2 RPM reaches 52 percent, the start fail time delay is activated. After 2 seconds the engine start VALVE light illuminates by a ground through the N2 engine speed card 52 percent switch. The EICAS level B message L(R) STARTER CUTOUT is then displayed after 5 seconds. This message inhibits all other caution and advisory messages for 20 seconds. If this occurs, position the engine ignition and start control switch to OFF, and if necessary remove pneumatic supply to the starter. Some operators procedures may require the affected engine to be shut/down. CAUTION: IF VALVE IS NOT CLOSED WHEN N2 INDICATION SHOWS 50% RPM, STARTER MAY BE DAMAGED.
HOLDING COIL
N2 52%
FULL OPEN STARTER CONTROL VALVE
SPEED CARD (P50)
ENG START 1 (P6) MD&T
A A
5 SEC
2 SEC ENG START VALVE (P10)
5 SEC ALL LEVEL B AND C MESSAGES INHIBITED FOR 20 SEC
1
GND IF NOT FULL CLOSED
2
GND IF NOT FULL OPEN
L(R) STARTER CUTOUT (B) L(R) ENG STARTER (C) PRIMARY ENGINE DISPLAY
EICAS
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ENGINE IGNITION SYSTEM - ENGINE IGNITION LEADS, PLUGS AND START CONTROL Location The ignition start control and select switches are located on the P-5 overhead panel in the engine ignition and start control panel.
Ignition Select Switch
Characteristics The conductor is 14 AWG stranded copper wire with silicone rubber insulation within a flexible conduit. The conduit contains an inner copper braid and an outer braid of nickel wire. Tubular plastic covers the cold section of the lead and an air cooling jacket covers the hot section. Fan air, used for cooling the lead enters through perforations at the forward end as the cable passes through a plenum. After cooling the lead, the air is discharged through a concentric port just above the coupling nut at the igniter plug.
There are two positions: • Single • Both The switch allows either circuit 1 or 2 to be selected by the EEC, or both circuits to be selected. This selection is for both engines.
Ignition/Start Switches There is a separate switch for each engine. The switches have five positions. These positions are • • • • •
GND AUTO OFF CONT FLT
The switch is detented in the AUTO position to prevent inadvertent selection of other switch positions.
Location The leads run from the exciter box location at the 7 o'clock position on the left fan case, to the igniter plugs on the compressor rear frame at the 3 and 4 o'clock position.
Igniter Plugs The igniter plug is a surface gap type used to ignite fuel within the combustion chamber. A coupling nut secures the igniter plug into a recessed adapter bolted into the compressor rear frame at two places, 4 and 3 o'clock. The immersion depth of the igniter plug is preset at the factory using spacers under the adapter. No depth check is required.
Safety Precautions Due to the high voltages, care should be taken with all ignition system components. See the following WARNING:
WARNING: IGNITION SYSTEM VOLTAGE IS DANGEROUSLY HIGH. IGNITION SWITCH MUST BE IN OFF POSITION BEFORE REMOVAL OF ANY IGNITION COMPONENTS. ALLOW SEVERAL MINUTES TO ELAPSE BETWEEN OPERATION OF IGNITION SYSTEM AND REMOVAL OF COMPONENTS. UPON DETACHING CABLE FROM IGNITER PLUG, DISCHARGE CURRENT BY GROUNDING CABLE TERMINAL TO ENSURE COMPLETE DISSIPATION OF ENERGY FROM THE SYSTEM. SEVERE INJURY COULD RESULT.
PNEUMATICS ENG START R
L SINGLE
BOTH
PNEUMATIC
VALVE
VALVE
STARTER GND
IGNITION/ START CONTROL SWITCHES
AUTO
OFF
GND
AUTO OFF
CONT
CONT
FLT
FLT
IGNITION SELECT SWITCH PNEUMATIC STARTER CONT VALVE
ENGINE IGNITION AND START CONT PANEL (P5)
IGNITION EXCITER 1
STARTING/ IGNITION
L
FUEL CONTROL
IGNITER PLUG 1 (4:00)
R
RUN
IGNITION EXCITER 2
CUTOFF
CHANNEL A CHANNEL B (SAME AS CHANNEL A)
EEC
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IGNITER
PLUG 2 (3:00)
FUEL CONTROL SWITCHES (P10)
B767-3S2F
ACCESSORY GEARBOX
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ENGINE IGNITION SYSTEM - IGNITION ELECTRICAL POWER SUPPLY SYSTEM Power Power is supplied to ignition exciter 1 from the 115 volt ac left main bus or alternately from the 115 volt ac standby bus. Power for ignition exciter 2 is supplied from the 115 volt ac right main bus or the 115 volt ac standby bus.
Ignition Select and Start Control Switches The engine ignition and start control panel located on the pilots' P-5 overhead panel contains the ignition select switch and the Ignition / Start switches for the left and right engines. The switch allows Single or Both exciters to be selected. The switch allows power to the exciters as follows: • GND: ignition is enabled for the EEC selected igniter • AUTO: ignition is enabled for the EEC selected igniter when thermal anti-ice is "ON" or if the slats are extended • OFF: no ignition • CONT: ignition is enabled continuously for the EEC selected igniter • FLT: ignition is enabled for both igniters bypassing the ignition select switch In all cases the Fuel / Ignition control relay must be de-energized to enable ignition. This requires the fire handles be in the NORM position and the fuel control switch in RUN.
Ignition Exciters The two independent exciter units are mounted on the engine fan case, left side at 7:00 o'clock. They are electrical capacitors that are enclosed in welded steel cases.
Control and Operation Ignition is controlled as a function of the ignition select switch, the ignition / start control switches, the fuel / ignition control relay, engine thermal anti-ice relay
and slat position. The EEC actually selects the ignition plug to fire in the SINGLE position. The EEC alternates Igniter plugs every other engine start in this position. In the BOTH position the EEC selects igniter plugs One and Two to fire together.
Displays and Indications If the left or right AC bus is unpowered, the associated power sense relay No. 1 allows standby bus power to the system. The power sense relay No. 2 provides a ground signal to EICAS. This causes the maintenance message IGN 1(2) STBY BUS to appear.
RUN
NORM
L AC BUS L ENG IGN 1 L AC BUS L ENG BUS PWR SEN
CUTOFF K158 (P11)
FUEL/IGN CONT 2 (P36)
FIRE
FUEL CONT SW (P10)
FIRE SW (P8)
L ENG IGNITER PLUG 1
28V DC BAT BUS L ENG FUEL CONT VALVE
L ENG IGN EXCITER 1
P11 L ENG IGNITER PLUG 2
CHANNEL A IGN SELECT LOGIC
R AC BUS L ENG IGN 2
L ENG IGN EXCITER 2
IGN SELECT LOGIC R AC BUS R ENG BUS PWR SEN
K608 (P11)
FUEL/IGN CONT 1 (P36)
CHANNEL B L ENGINE EEC
ECS/MSG
IGN 2 STBY BUS IGN 1 STBY BUS
STBY BUS STBY IGN 1
CHANNEL A
L AC BUS R ENG IGN 1
IGN SELECT LOGIC K607 (P11)
STBY BUS STBY IGN 2
FUEL/IGN CONT 2 (P36)
R ENGINE EEC
R AC BUS R ENG IGN 2
CUTOFF K159 (P11)
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R ENG IGN EXCITER 2
CHANNEL B
RUN
FUEL/IGN CONT 1 (P36)
R ENG IGNITER PLUG 2
EICAS
IGN SELECT LOGIC
FUEL CONT SW (P10)
R ENG IGN EXCITER 1
NORM
FIRE FIRE SW (P8)
IGNITION SYSTEM POWER TRAINING MANUAL FOR TRAINING PURPOSES ONLY
28V DC BAT BUS R ENG FUEL CONT VALVE P11
R ENG IGNITER PLUG 1
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THRUST REVERSER - THRUST REVERSER SYSTEM 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.
FLEXIBLE DRIVE SHAFT
FAN EXHAUST
TURBINE EXHAUST ANGLE GEARBOX BALLSCREW ACTUATOR (2)
FORWARD THRUST CONFIGURATION
TRANSLATING COWL CDU/ ACTUATOR THRUST REVERSER - STOWED
FAN EXHAUST OUTER FAN DUCT
DEFLECTORS
INNER FAN DUCT
TURBINE EXHAUST BLOCKER DOOR DRAG LINKS (6) BLOCKER DOORS (6)
REVERSE THRUST CONFIGURATION
FAN DUCT
THRUST REVERSER SYSTEM B767-3S2F Page - 177
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THRUST REVERSER - DEPLOYED
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THRUST REVERSER - 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 aluminum. 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. 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. 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. CAUTION: DO NOT OPERATE ENGINE IN REVERSE THRUST WITH DEFLECTORS MISSING. DAMAGE TO THE REVERSER MAY RESULT.
SPRING RETAINER CLIP
SPRING (4)
BLOCKER DOOR
DRAG LINK BLOCK
DEFLECTOR QTY TYPE (R ENG) A B C D E F G
INNER FAN DUCT BLOCKER DOOR DRAG LINK
INNER FAN DUCT COWL HINGE
15 3 1 2 2 4 5
DESCRIPTION RADIAL -43 |5 FWD SKEWED -25 |5 FWD, 45 LH SKEWED -25 |5 FWD, 45 RH BLANK CURVED STRONGBACK -45 |5 FWD, LH CURVED STRONGBACK -45 |5 FWD, RH SPOILED RADIAL 0-10 FWD
TRANSLATING COWL
FWD LINK SUPPORT
LINK PIN
DEFLECTOR BOLT
TRIWING SCREW
GANG CHANNEL
THRUST REVERSER FIXED STRUCTURE
A A A 1 A 32 A 2 31 A 3 30 A A 4 29 A C 5 28 A F 6 27 F 7 26 A AFT LOOKING FORWARD 25 E F 8 24 E F 9 INBD 23 A G 10 11 22 G A 21 12 G 13 20 A 19 A G 14 17 18 B G 15 16 B D D B RIGHT ENGINE SHOWN - LEFT ENGINE SIMILAR
DEFLECTOR INSTALLATION (TYPICAL)
THRUST REVERSER - DEFLECTORS B767-3S2F Page - 179
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THRUST REVERSER DEFLECTORS
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THRUST REVERSER - THRUST REVERSER SYSTEM OPERATION General Thrust reversers are used by the flight crew to decelerate the airplane immediately after landing or during a refused takeoff. 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 be available. Deploy When reverser deployment is commanded, switch and relay logic provide power to unlock the electro-mechanical brake, to energize the directional pilot valve and to open the Thrust Reverser Pressure Regulating and Shut Off Valve (T/R PRSOV). Air from the T/R PRSOV flows to the left and right CDUs and to the DPV. An air signal from the DPV to the CDU arms the CDU to the deploy mode. Air motors in the CDUs drive ballscrew actuators attached to 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 CDUs. The air motors in the CDUs 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. Electronic position feedback on each half of the thrust reverser, provided to the EEC allows the throttle interlock solenoid to operate. The crew can then move the reverse thrust levers to the high power position.
Engine Operation During the approach to landing, the engine is not permitted to 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 discrete’s 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. 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 in-flight 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.
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.
Stow When the thrust reverser is commanded to stow, air from the T/R PRSOV flows to the left and right CDUs and the DPV. Now the DPV remains closed, blocking the air signal to the CDUs. This arms the CDUs 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 brakes. The system is now locked in the stowed position by the CDU cone brakes and by the electromechanical brakes.
28V DC
AIR/GND SWITCH
28V DC
FIRE SWITCH
28V DC
FIRE SWITCH
T/R CONTROL SWITCH
10 DEGREE SW
ITCH
T/R DPV SWITCH
29 DEGREE SWITCH
AIR/GND RELAY
T/R AUTOSTOW LOCK SWITCH
AIR/GND RELAY
RELAY LOGIC THRU SEQUENCING RELAY K2184 AND TRAS LOCK RELEASE RELAY K2188
ELECTROMECHANICAL DISK BRAKE
G 29 D E
W REE S
ITCH
REV ISLN P10
REV 26.1
10 6
CDU POSITION SWITCH MODULE
2
PRESSURE SWITCH
T/R PRSOV PRSOV BLEED AIR
N1
EICAS DISPLAY
L(R) REV ISLN VAL (C)
INDICATION LOGIC AND BITE RELAY MODULE
L(R) REV ISLN VAL (M)
EICAS YELLOW GREEN CDU POSITION FEEDBACK TRANSDUCER
DIRECTIONAL PILOT VALVE (DPV) CENTER DRIVE UNIT (CDU) FLEXIBLE DRIVE SHAFT
ANGLE GEARBOX AND BALLSCREW ACTUATOR
TRANSLATING COWL
THRUST REVERSER OPERATION B767-3S2F Page - 181
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EEC
INTERLOCK ACTUATOR
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THRUST REVERSER 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 pilot's 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 micro-switch 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 degrees. The DPV control switch closes when the reverser thrust lever is raised above 29 degrees. This signals the directional pilot valve to open, directing air to the DEPLOY side of the CDU air motor. At 29 degrees the TRAS lock switch closes, providing power to several relays which unlock the electromechanical brakes and signal the T/R PRSOV to open.
REVERSE THRUST LEVER
FWD
T/R CONTROL SWITCH (OPERATES AT 10 DEG)
REVERSE THRUST LEVER
FORWARD DRUM (REF)
FORWARD THRUST LEVER
T/R TRAS LOCK SWITCH (OPERATES AT 29 DEG)
T/R DIRECTIONAL PILOT VALVE SWITCH (OPERATES AT 29 DEG)
T/R CONTROL SWITCH COVERS
THRUST REVERSER CONTROL SWITCHES B767-3S2F Page - 183
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THRUST REVERSER - ELECTRO-MECHANICAL (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.
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 28VDC 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.
ELECTRICAL CONNECTOR
MANUAL RELEASE HANDLE ELECTRO-MECHANICAL BRAKE
BRACKET FLEXSHAFT ANGLE GEARBOX DRIVE PAD
CENTER DRIVE UNIT
ELECTRO-MECHANICAL (TRAS) BRAKE B767-3S2F Page - 185
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THRUST REVERSER PRESSURE REGULATING VALVE Pressure Regulating and Shutoff Valve (T/R PRSOV) 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 pre-cooler. 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.
T/R PRSOV Operation 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 modulates 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.
SOLENOID VALVE
PRESSURE REGULATOR (70 PSI)
RELIEF VALVE (150 PSI)
SOLENOID
PNEUMATIC ACTUATOR
INLET
POPPET VALVE
INLET
OUTLET
THRUST REVERSER PRESSURE REGULATING VALVE B767-3S2F Page - 187
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OUTLET
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DIRECTIONAL PILOT VALVE General The directional pilot valve (DPV) is a solenoid controlled pressure operated valve. Switch and relay logic control the solenoid. Air pressure is supplied when the T/R PRSOV is open. When the DPV is open, it provides air pressure to both halves of the thrust reverser for that engine. This air pressure, called signal air, operates on a piston within each of the CDUs. The result of the piston motion is to change the position of the directional control valve (DCV) in each CDU. The main flow of air from the T/R PRSOV into the air motor is determined by the position of the DCV. The air motor direction of rotation is reversed as the position of the DCV is changed. One direction of motor rotation moves the sleeves to the deployed position. The opposite direction of air motor rotation moves the sleeves to the stow position. The operation of the air motor and the DCV is discussed later. 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 on each 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 twoposition microswitch.
Operation When reverse thrust is selected, the solenoid is energized and the ball valve moves down and closes the vent. The poppet valve opens to let air pressure from the T/R PRSOV go to the directional control valve. When the thrust reverser system is in the stow position, the solenoid is deenergized. Air pressure from the T/R PRSOV is blocked. The signal air lines to both CDU directional control valves are vented through the DPV ball valve to ambient. 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.
SOLENOID AMBIENT VENT BALL VALVE ASSEMBLY
PRESSURE SWITCH
OUTLET TO CDU DCV
POPPET VALVE DPV FILTER
FWD
THRUST REVERSER TORQUE BOX
PRESSURE SWITCH
SOLENOID
DPV VALVE BODY
AIR IN FROM T/R PRSOV
OUTLET TO RIGHT CDU DCV
INLET
DPV FILTER OUTLET TO LEFT CDU DCV
DIRECTIONAL PILOT VALVE B767-3S2F Page - 189
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THRUST REVERSER - 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 rod assembly. The gearbox has two flexible drive shaft output drives and a manual drive pad. A Directional Control Valve (DCV) includes a directional valve, a helix rod and spring, and a valve actuator piston. The DCV is spring-loaded 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.
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 the translating cowl to the deploy position. 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.
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.
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. CAUTION: DO NOT REMOVE CLEVIS PIN RETAINING CLIP BOLT. BACK BOLT OUT ENOUGH TO ROTATE RETAINING CLIP. REMOVAL OF BOLT WILL DAMAGE NUTPLATE. Disconnect the feedback cable and the rotary flexible drive shafts. 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. Note:
Be sure to reference the aircraft M/M when ever you perform any maintenance operation.
SIGNAL AIR FROM DPV
CDU POSITION FEEDBACK TRANSDUCER MANUAL BRAKE RELEASE HANDLE
STOW VALVE ACTUATOR PISTON
IN DEPLOY
AIR MOTOR
HELIX ROD
PIVOTED LEVER
DIRECTIONAL CONTROL VALVE (STOW POSITION)
ACTUATOR (CONE) BRAKE DCV (NEUTRAL POSITION)
AIR MOTOR
AIR INLET
TO POSITION INDICATING SWITCHES OVERRIDE LINKAGE
DEPLOY STOP
STOW STOP
STOP ROD
MECHANICAL SWITCH INPUT STOP ROD
BALLNUT
CDU POSITION INDICATING SWITCH MODULE
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GEARBOX LOCKOUT PLATE
CENTER DRIVE UNIT OPERATION Page - 191
BALLSCREW
ELECTRICAL CONNECTOR (BACK SIDE)
TO DCV
B767-3S2F
STOP ROD
TORQUE TUBE
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SQUARE DRIVE
B767-3S2F Page - 192
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THRUST REVERSER - 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 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. CAUTION: ENSURE THAT THE DRIVE SHAFT CORE DOES NOT SLIDE OUT OF OUTER CASE WHEN REMOVING THE ROTARY FLEXIBLE DRIVE SHAFT. CAUTION: 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. Note:
When installing a gearbox and actuator the side plate on the gearbox must be facing inward.
SPLINED OUTPUT DRIVE
BALLNUT
ACTUATOR TUBE
BALLSCREW
ROD END BEARING
BALLSCREW ACTUATOR
ANGLE GEARBOX
THRUST REVERSER TORQUE BOX RETAINING CLIP AND BOLT
FLEXIBLE DRIVE SHAFT
ANGLE GEARBOX AND BALLSCREW ACTUATOR
CLEVIS PIN
FACEPLATE (INWARD FACING NOT SHOWN) TRANSLATING COWL
FWD
CAPPED END
ANGLE GEARBOX AND BALLSCREW ACTUATOR B767-3S2F Page - 193
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THRUST REVERSER - ELECTRICAL OPERATION 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: • T/R control switch (S5) • T/R DPV control switch (S11) • TRAS lock switch (S21). The T/R control switch (S5) is the first to close at approximately 10 degrees of reverse thrust lever rotation. At approximately 29 degrees 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 CDUs. There is approximately a 160 millisecond window between the TRAS brake release and the CDUs 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 CDUs, translates to the fully deployed position.
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 CDUs. Two seconds after removal of the pneumatic operating pressure from the thrust reverser system, the 28 VDC power is removed from the electro-mechanical brake solenoids and the brakes engage again.
AIR 28V DC STBY BUS C1491 L ENG T/R CONT ALT (11D5)
POWER TO COIL OF K10234 L ENG T/R DISAGREE NORM
GND
STOWED
DEPLOY
L T/R DPV (L ENG)
S11 (29 DEG) L T/R DPV CONT
K895 SYS 1 AIR/GND (P36)
UNSTOW NOT DEPLOY
C1482 L ENG T/R CONT (11L6)
28V DC L BUS
STOW
FIRE AIR
STOW
DEPLOYED
S37 L ENG FIRE SWITCH (P8)
C1487 28V DC L BUS PWR SENSE (11M3)
DEPLOY S5(10 DEG) L T/R CONT
K897 28V DC L BUS PWR SENSE (P11)
GND
RH T/R LOGIC SW (R CDU-L ENGINE)
K895 SYS 1 AIR/GND (P36)
NOT DEPLOY
POWER TO COIL OF K1025 L T/R DEPLOY IDLE (APP IDLE CMD)
NOT DEPLOY 28V DC STBY BUS C1576 L ENG T/R TRAS LK CONTROL (11D18) P11 CB PANEL ASSY FAULT LATCHING GND FOR TRRM L ENG RESTOW COMMAND LAMP
DEPLOYED STOWED
UNSTOW FULLY DEPLOYED
LH T/R LOGIC SW (R CDU-L ENGINE) STOWED
K1023 L T/R DEPLOY (P36)
STOWED OR DEPLOY
POWER TO COIL OF K1034 L T/R VALVE RELAY (APP IDLE CMD)
ONE SEC T/D ON RELEASE AFTER STOWED
POWER TO COIL OF K1021 L T/R PNEU VLV
V360 L ENG T/R PRSOV (L STRUT)
UNLATCH 2 SEC
DEPLOY T/D
UNSTOW OR NOT DEPLOY K26 RLY-L T/R UNSTOW (P36)
TD-L TRAS UNLATCH (P33)
K2184 RLY-L T/R SEQ (P33)
UNSTOW K2182 L TRAS LK REL (P33)
R702 DIO - LH TRAS LK RLY (P33) POWER TO COIL OF K2186 L T/R TRAS UNLK LH TRAS SOL
AIR
STOW
GND
DEPLOY
K2157 RLY AIR/GND (P37)
S21 (29 DEG) LEFT TRAS LK
RH TRAS SOL M LATCH K2188 L TRAS LK REL CONT (P33)
THRUST REVERSER ELECTRICAL OPERATION B767-3S2F Page - 195
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R704 DIO - RH TRAS LK RLY (P33)
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THRUST REVERSER - 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.
T/R Malfunction Indications After the airplane has been on the ground for 60 seconds, 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 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 inflight fault in the reverser system
Thrust Reverser Relay Module (TRRM) 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 enable 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. The thrust reverser relay module will latch a fault if 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 PROSOV as indicated by the pressure switch mounted on the directional pilot valve. The LED labeled PRSOV PRESSURE will be illuminated.
REV (YELLOW REV (GREEN)
MUX
L T/R IN TRANSIT L T/R DEPLOYED
NVM LATCH
L REV ISLN VLV (LEVEL M)
L REV ISLN VLV (LEVEL C)
2 SEC
STOWED
AIR
UNSTOW SOFTWARE
GND
NOT DEPLOY
K2175 AIR/GND SYS 1 (P36)
FAULT LOGIC GND REQ'D SOFTWARE
TEST S1 L TEST ENABLE L3 L ENG RESTOW COMMAND
DEPLOYED R T/R LOGIC SW (R CDU-L ENG)
L OR R EICAS COMPUTER (E6)
R3 UNLATCH
RESET S3 L RESET CR3
A
A
MD&TR10117 L T/R IND (P37)
L5 REV ISLN FUEL CONT PNL (P10)
LATCH
FAULT FAULT
5 SEC
LATCH NOT DEPLOY
M10440 L T/R ISN VLV DELAY (P36)
P11 CB PNL ASSY
K7 L T/R FAULT LATCH
UNSTOW BYPASS OF K26 TO KEEP TRAS LK SOLENOIDS ENERGIZED DURING STOW
28V DC STBY BUS C1480 L ENG T/R IND (11D13)
K3 L ENG RESTOW COMMAND LATCH
STOWED
UNLATCH
NOT DEPLOY CR1 DEPLOYED R1
L T/R LOGIC SW (L CDU-L ENG)
NORMAL K10358 L T/R ISLN DET (P33)
L1 L ENG TRAS UNLOCK
DEPLOY
60 SEC
K1023 L T/R DEPLOY (P36)
AIR
T / D
LATCH K1 L ENG TRAS UNLOCK LATCH
NORMAL K9 L ENG FAULT DET
GND
STOW
K11 L T/R FAULT DELAY
K2175 AIR/GND SYS 1 (P36) UNSTOW
M3 TD L T/R RESTOW COMMAND
K26 L T/R UNSTOW (P36)
LAMP TEST
LOCK LEFT ENG
RIGHT ENG
AIR
RESTOW PRESSURE
5 SEC
STOWED OR DEPLOYED
PSEU RESTOW
M1 TD L ENG TRAS UNLOCK
AIR
L5 L ENG PRSOV PRESS
5 PSI
RESET
LATCH
M7 TD L T/R FAULT
GND PRESS AND HOLD FOR TEN (10)SECONDS
K178 SYS 1 AIR/GND (P36)
GND
NOT DEPLOYED
K10234 L ENG T/R DISAGREE (P36)
FIRE SWITCH NORMAL AND ON GROUND
K1021 L T/R PNEU VLV (P36)
UNLOCK K2186 L T/R TRAS UNLOCK (P33)
TRAS SOLENOID POWER IS APPLIED
PNEUMATICS COMMANDED FOR DEPLOY
CR5
PRSOV PRESS SWITCH (LEFT STRUT)
R700 L PRSOV PRESS (P33)
TEST S5 LAMP TEST 5 SEC M5 TD L ENG PRSOV PRESS THRUST REVERSER RELAY MODULE (E2-6 OR E1-4)
THRUST REVERSER INDICATING SYSTEM OPERATION B767-3S2F Page - 197
R5
5 SEC
HIV PRESSURE
GROUND MODE TEST ENABLE
UNLATCH
5 SEC
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PC CARD
K5 L ENG PRSOV PRESS LATCH
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THRUST REVERSER - TRANSLATING COWL MANUAL DEPLOY/STOW General This procedure covers manual cowl translation (deploy or stow) of the translating cowl using either a manual speed wrench or an air-powered wrench. Each cowl is operated independently of the other using this procedure. Do not extend either translating cowl if the thrust reverser is opened more than the 34degree (first stick) position.
WARNING: YOU MUST CAREFULLY FOLLOW THE INSTRUCTIONS IN THIS TASK. IF YOU DO NOT, THE THRUST REVERSER CAN ACCIDENTLY OPERATE AND CAUSE INJURY TO PERSONS AND DAMAGE TO EQUIPMENT.
Deploy Open the applicable circuit breakers on the P11 panel to remove power from the thrust reverser actuation system; install DO-NOT-CLOSE identifiers on the circuit breakers. Deactivate the spoiler/speed brake control system. Insure that the reverse thrust levers are fully forward, and attach a DO-NOT-OPERATE tag. Make sure that a pneumatic source is not connected to the thrust reverser. Open the fan cowl panels. Make sure that the D-shaped pressure relief door is closed and latched. If the thrust reverser is opened to the 34-degree (first stick) position, make sure that the leading edge slats are fully retracted. Pull up on the manual release handle to unlock the electro-mechanical (TRAS lock) brake. Pull the cone brake release handle out and away from the CDU until the detent is felt. Remove the two bolts that attach the lockout plate to the manual drive pad on the bottom of the CDU. Put a 1/4-inch square-drive into the CDU manual drive. Turn the square-drive on the CDU to extend the translating cowl. Less than 10 pound-inches of torque should be applied. Open the other thrust reverser half if it is necessary.
WARNING: DO THE DEACTIVATION PROCEDURE FOR THE SPOILER/ SPEED BRAKE SYSTEM OR REMOVE ALL PERSONS AND EQUIPMENT AWAY FROM THE SPOILERS. THE SPOILERS CAN RETRACT QUICKLY AND CAUSE INJURIES TO PERSONS AND DAMAGE TO EQUIPMENT. (REF AMM 2761-00/201) CAUTION: DO NOT OPEN THE THRUST REVERSER HALF TO MORE THAN THE 34-DEGREE (FIRST STICK) POSITION IF THE TRANSLATING COWL IS EXTENDED. DAMAGE TO THE TRANSLATING COWL OR THE STRUT CAN OCCUR.
CAUTION: IF YOU USE AN AIR WRENCH TO EXTEND/RETRACT THE TRANSLATING COWL, LOOK FOR MOVEMENT OF THE FEEDBACK ROD WHEN THE TRANSLATING COWL IS ALMOST FULLY EXTENDED/RETRACT. WHEN YOU SEE MOVEMENT, REMOVE THE AIR WRENCH AND FULLY EXTEND/RETRACT THE TRANSLATING COWL WITH A MANUAL WRENCH. THE CDU WILL LOCK IF THE STOPS ARE ENGAGED, AND DAMAGE TO THE CDU CAN OCCUR.
Stow CAUTION: MAKE SURE THAT THERE IS NO EQUIPMENT IN THE AREA AFT OF THE THRUST REVERSER. DAMAGE CAN OCCUR IF THE THRUST REVERSER HITS THE EQUIPMENT. CAUTION: WHEN YOU MANUALLY MOVE THE THRUST REVERSER, LOOK FOR THE TOP AND BOTTOM BALLSCREW ACTUATORS TO TURN. IF YOU DO NOT SEE THESE BALLSCREW ACTUATORS TURN, DO A CHECK FOR FLEXSHAFTS THAT ARE BROKEN OR GONE.
Prepare the thrust reverser for stowing the thrust reverser as you did for deploying the translating cowl. Put a 1/4-inch square-drive into the CDU manual drive. Turn the square-drive on the CDU to retract the translating cowl. Less than 10 pound-inches of torque should be applied. When the translating cowl is about one inch from the fully retracted position, push the stow rig button on the CDU. Stop turning the CDU when the stow rig pin moves and then starts to move out again. Turn the wrench in the direction that aligns the rig pin plunger with the groove in the CDU actuator. Measure to make sure that the clearance between the torque box and the translating cowl is between 0.0600.150 inch (1.5-3.8 mm). Restore the airplane to normal.
MANUAL BRAKE RELEASE HANDLE
CDU STOW RIG INDICATOR BUTTON MANUAL RELEASE HANDLE
RIG INDICATOR PLUNGER GROOVE
CDU RIG WINDOW
CDU CONNECTOR
CDU TRANSLATING COWL SPRING TORQUE WASHER BOX
MANUAL DRIVE PAD LOCKOUT PLATE
CDU POSITION SWITCH MODULE
TRANSLATING COWL MANUAL DEPLOY / STOW B767-3S2F Page - 199
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THRUST REVERSER - TRANSLATING COWL POWER DEPLOY/STOW USING AIR APPLIED 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-degree (first stick) position.
WARNING: BE SURE TO COMPLY WITH ALL MM WARNINGS, CAUTIONS AND ADVISORIES. FAILURE TO DO SO MAY RESULT IN PERSONAL INJURY OR DAMAGE TO EQUIPMENT.
Deploy Open the selected circuit breakers on the P11 panel and install DO-NOTCLOSE identifiers. (see MM) Deactivate the spoiler/speedbrake control system, ensure the reverse thrust levers are in the forward (stow) position, and ensure that the thrust reverser is not open beyond the 34-degree 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. Place the reverse thrust levers to the reverse idle position and allow translating cowl to fully deploy.
Stow Provide pneumatic power and place the 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.
MANUAL RELEASE HANDLE
ENGINE AIR SUPPLY
BLUE CAP COVERING GROUND CONNECTION
TRANSLATING COWL TORQUE BOX CENTER DRIVE UNIT
TRANSLATING COWL POWER DEPLOY / STOW SUPPLYING AIR THROUGH CDU B767-3S2F Page - 201
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THRUST REVERSER - TRANSLATING COWL DEPLOY / STOW WITH GROUND SERVICE SWITCH General This procedure covers power translation of the thrust reverser translating sleeve using air from the opposite engine, external pneumatics connection or the APU. This air in the pneumatic system normally can not back-flow through the Engine PRSOV to the T/R PRSOV. This process electronically opens the PRSOV using the ground service switch. This is a guarded switch, spring loaded to the “OFF” position that is located next to the engine oil tank.
WARNING: BE SURE TO COMPLY WITH ALL M/M WARNINGS, CAUTIONS, AND ADVISORIES. FAILURE TO DO SO MAY RESULT IN PERSONAL INJURY OR DAMAGE TO EQUIPMENT. Refer to the applicable MM for Spoiler / Speedbrake deactivation. Inadvertent spoiler movement caused by actuating thrust levers could result in serious injury to personnel. Ensure reverse thrust levers are in the forward thrust (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 reversers will move when the T/R lever is moved to the reverse thrust position. Ensure area aft of the T/R is clear of personnel and equipment before operating the thrust reverser. Note:
With pneumatic power provided, a deployed thrust reverser will stow if electrical power is lost to the directional pilot valve.
WARNING: WHEN MAINTENANCE IS PERFORMED ON OR NEAR THE T/R THE SYSTEM SHOULD BE LOCKED OUT PER THE MM. Deploy Open the selected T/R circuit breakers on the P11 panel and install “DO NOT CLOSE” identifiers. Deactivate the spoiler speed brakes. Ensure the thrust reverser levers are in the forward thrust position (stowed). Ensure T/R is not open beyond the 34 degree position, and that the core cowl panels are removed or closed. Open the fan cowl panels. Provide pneumatic power to the airplane per MM. Push the applicable L or R ENG OFF 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. Place the reverse thrust levers to the reverse idle position. Lift the guard on the PRSOV ground service switch. Push the switch to the up position and hold it. Allow the translating cowls to fully deploy before releasing the switch.
Stow Provide pneumatic power. Push the applicable “L or R ENG OFF” switch light on the air supply module on the P5 panel to the open position and place the reverse thrust lever to the forward position (stowed). Lift the guard on the PRSOV ground service switch and push the switch up. Hold the switch until the T/R is fully stowed. Release the ground service switch. Ensure the T/R is fully stowed by checking the gap between the torque box and the translating cowl is between 0.060 and 0.150 inch at the CDU. Return the aircraft to normal configuration.
ENGINE OFF SWITCH-LIGHTS
DIRECTIONAL PILOT VALVE ELECTRO-MECHANICAL BRAKE (TRAS LOCK)
TO OTHER CDU
FROM PNEUMATIC SOURCE
REVERSE FLOW SOLENOID (PULL TYPE)
BLEED DL UE CA TK
OVHT
a
APU
L ENG O F F
OVHT
ADP
R ENG
V A L V E
BLEED AIR SUPPLY PANEL (P5)
PRSOV
ELECTRICAL HEX FOR CONNECTOR MANUAL OPERATION
CDU
THRUST REVERSER PRESSURE REGULATING AND SHUTOFF VALVE (T/R PRSOV)
PNEUMATICS
B767-3S2F Page - 203
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OIL TANK (REF)
TRANSLATING COWL POWER DEPLOY / STOW WITH GROUND SERVICE SWITCH ATA 78-00 EFF - ALL
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O F F
GROUND SERVICE SWITCH
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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.
Install three red deactivation plates Install both lockout plates on the CDU drive pad Verify T/R position on EICAS Close fan cowls Reset pulled CB’s Pull out and collar effected CB’s
CAUTION: DAMAGED OR BROKEN DRAG LINKS MUST BE REMOVED. ANY EFFECTED BLOCKER DOORS MUST BE TAPED SHUT.
Deactivation CAUTION: 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. CAUTION: THIS PROCEDURE IS FOR GROUND INADVERTENT THRUST REVERSER TRANSLATION MAY OCCUR IF PROCEDURE IS USED TO DEACTIVATE THRUST REVERSER FOR FLIGHT DISPATCH. 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 CDUs and attach REVERSER DEACTIVATED pennants.
Lockout Note:
• • • • • •
When locking out a Thrust Reverser for dispatch be sure to reference the MEL for specific instructions. Lockout and test instructions must be complied with prior to aircraft dispatch.
The following steps are required to be performed to lockout a Thrust Reverser (T/R) for flight dispatch: • Remove the lockout plate from the CDU manual drive pad • Check the running torque of the T/R system (
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