51695947 B 757 General Familiarisation in Maintenance

5/24/2018 51695947 B 757 General Familiarisation in Maintenance

1/167

BOEING 757BOEING 757BOEING 757BOEING 757BOEING 757A PDF STA PDF STA PDF STA PDF STA PDF STORYORYORYORYORY

DIGITAL VERSION OF BOOK PRODUCED BY LORENZO SOLBERGHE & ZDENKO SIMAC

Techtraining


2/167

SECTION TITLE ATA

1 Introduction 6, 9, 12

2 Structures 51, 55, 57

3 Equipment Centers 11, 20, 23, 394 Flight Deck 11, 25, 33, 35, 39

5 EICAS 31

6 Electrical Power 24

7 Fuel System 28

8 Auxiliary Power Unit 49

9 Power Plant Rolls-Royce 71-80

10 Power Plant Pratt&Whitney 71-80

11 Hydraulics 29

12 Landing Gear 32

13 Flight Controls 27

14 Environmental Systems 21, 36

15 Ice and Rain Protection 30

16 Fire Protection 26

17 Cabin Systems and Lighting 25, 33, 35,

38, 52, 5618 Cargo Systems 25, 33, 52

19 Communications 23

20 Indicating and Recording 31

21 Navigation 34

22 Autoflight 22

Glossary

Table of ContentsPa


3/167

StructuStructuStructuStructuStructu

Features

BASIC STRUCTURALDESCRIPTION

The 757 is metal low-wing mono-plane with full cantilever wing and

tail surfaces, semimonocoquefuselage, and fully retractablelanding gear. Its two power plantsare located on the wings on struts.

DESIGN PHILOSOPHY

Redundant structural load pathsand scheduled aircraft inspectionsare part of the damage-tolerantdesign philosophy. Fatigue testing,monitoring of high-time airplanes,

and continuing coordination be-tween Boeing and the airlinescomplete this design philosophy.

COMPOSITE MATERIALS

For high strength and stiffnesswith minimal weight, the 757incorporates substantial amountsof carbon, aramid, or fiberglasscomposite materials.

DESIGN SERVICE OBJECTIVE

Structure is designed to meetservice objectives in flight cycleswhich are typically achieved aftermore than 25 years of service.

CORROSION PREVENTION

The 757 uses the most advancedcorrosion prevention methodsavailable and meets or exceedsthe Inter national Air Transport

Association guidelines. Corrosionprevention systems are continuallyupdated to reflect the latest tech-nology and in service experience,ensuring a structurally superiorairplane.

757 AND 767 COMPARISON

Structures for the 757 were de-signed for ultimate strength, dam-age tolerance, ``and durabilityusing the same design philosophyused for the 767. The certificationbasis for the 757 is identical to

that for the 767.

Differences in the actual designload levels for major structuralcomponents of the airframe aresignificant, reflecting the differ-ences in configuration, size, andweight. Although actual designload levels may differ betweenmodels, structural efficiency hasbeen maintained by using similardesign working stress levels.

Structures for the 757 and 767 aremanufactured using basically thesame methods, materials, andfasteners. Exceptions are made toprovide the most effective struc-ture in terms of weight, cost, andairplane performance. For ex-ample, the main landing gearbeam is titanium on the 757 andaluminum on the 767.

Both airplanes use advanced

composites extensively in similarapplications such as control sur-faces (carbon) and secondaryfairings (carbon, aramid, fiberglasshybrids).

Fuselage Reference Diagra

Fuselage Materials

Wing

Wing Center Section

Horizontal Stabilizer

Vertical Stabilizer

EAT doc 51 Issue 1

Revision 00 date 01/06 Pa


4/167


Fuselage Reference Diagram

The fuselage is divided into bodyreference planes (see page 1-2).

The fuselage is also divided intoproduction manufacturing sections.

Section 41 (STA 159 to 439), theforward fuselage, contains theradome and forward pressurebulkhead, forward access door,nose gear wheel well, and forwardentry doors.

Section 43 (STA 439 to 900)contains the electrical equipment

bay access, forward cargo com-partment, and entry doors.

Section 44 (STA 900 to 1180)includes the mid fuselage, emer-gency exits (overwing exit model),pressure deck, and main gearbulkhead.Section 46 (STA 1180 to 1720)contains the emergency exits (fourdoor model), aft cargo compart-ment, and aft entry doors.

Section 48 (STA 1720 to 2005)includes the aft pressure bulkhead;auxiliary power unit (APU); control,service, and APU doors; and

horizontal and ver tical stabilizers.

Fuselage

The fuselage is a semimonocoquestructure primarily constructedfrom conventional 2024 and 7075aluminum alloys. Improved higherstrength materials for forgings andextrusions are used on keel beamand major body frame structure.External clad skins are reinforcedby longitudinal stringers and

circumferential frames on a20-inch (51-centimeter) spacing. Atypical cross section through thefuselage consists of an uppercircular lobe and a lower oval lobethat intersect at the passengerfloor level. Transverse floorbeamsare located at this intersection andare suppor ted by the frames. Thefuselage is designed to withstandinternal pressure and externallyapplied loads from flight and

ground operating conditions.

The radome forward of the forpressure bulkhead is hinged a

top and made of fiberglass skand honeycomb.

The flight deck has three windshields on each side, numberesequentially from forward to aThe No. 1 windshields are flatset into forged titanium framesThe No. 2 and 3 windshields acurved and set into forged alunum frames. The No. 2 windshcan be opened. Passenger win

dows are made frommoisture-resistant acrylic matand are mounted in one-piecealuminum forgings.

Fuselage Reference Diagram

EAT doc 51 Issue 1



5/167


Fuselage skins are manufacturedby chemical milling or machiningon the inside surface to providereinforcement at stringer locations,

cutouts, and splice areas. In theupper lobe, tear straps and cutoutdoublers are hot bonded to skins.Frames are used to maintain thefuselage cross section shape andto transfer loads into the skin.

Primary bulkheads of the fuselageare the forward pressure nose andmain gear wheelwells and the frontspar and rear spar, main landinggear, aft pressure, and horizontalstabilizer pivot bulkheads.

The passenger floor structure is abuilt-up grid system consisting offloorbeams, stabilizing straps, seator freight tie-down tracks, and floorpanels. It extends from the forwardpressure bulkhead at STA. 192 tothe rear pressure bulkhead at STA1720.

Seat tracks are fabricated fromaluminum extrusions and designedto allow placement of seats any-where along the floor. Galleys and

lavatories are attached to the floorstructure using special fittings.Special tracks made from stainlesssteel may be used to mount gal-leys when quick removal andreplacement of the galley is re-quired. Floor panels are lightweightlaminations composed of fiber-glass skins with an aluminum oraramid (Nomex) honeycomb core.

Cutouts in the fuselage for pas-senger and cargo doors and elec-

trical/ electronics access arereinforced.

The passenger and lower-lobecargo doors are plug-type designsthat are not load carrying and thatact as simple pressure plugs. Themain deck cargo door on thefreighter model is outward openingand carries fuselage loads.

All aluminum fuselage parts aanodized or alodined and primwith corrosion-inhibiting primedetail. In addition, detail parts

located beneath the passengefloor receive a coat of whiteenamel. Fuselage parts in thelower lobe that are in contact the skin or are on the exteriorsealed on contact surfaces.Water-displacing corrosion inhtors are applied to the interiorfuselage structure and to seleareas of the exterior after allfinishing and sealing. The lowelobe uses a drainage systemconsisting of drain holes and p

through the structure to permiliquids to reach numerous extenally serviceable pressurized valves mounted on the bottomcenterline of the fuselage.

757 Structures

EAT doc 51 Issue 1



6/167


Materials

COMPOSITE MATERIALS

The airplane structure incorpo-

rates advanced composite materi-als for their high strength-to-weightratio.Significant weight savings havebeen made by substituting carbonand aramid advanced fiber com-posite materials for conventionalmetal and fiberglass construction.These materials also provideimproved fatigue, corrosion, andsonic resistance and superioraerodynamic surfaces.

Carbon fiber is used for the pri-mary movable surfaces such asthe ailerons, elevators, r udder,spoilers, and aft flaps.Carbon-reinforcedaramid-fiberglass hybrids are usedfor secondary fairing structures.Carbon is used in both the wovenfabric polyform and unidirectionalfiberply tape forms. Aramid andfiberglass plies are only used inthe woven fabric form.

High-st rength 350F (177C) curecarbon-epoxy pre-impregnated(prepreg) raw material is used forthe majority of the components,

and 250F (121 C) cure prepregfor the majority of the secondaryhybrid components. Large surfacepanels use honeycomb sandwichconstruction with solid laminateedge bands for attachment tosupporting structure. Nonmetallicaramid honeycomb core is sur-faced with composite face skinstailored to provide minimum weightmaximum stiffness components.To prevent galvanic corrosion toaluminum components in contact

with carbon materials, specialprotective systems are used.Fiberglass or aramid plies areco-cured to the carbon contactsurface.

Each aluminum component isanodized, primed, and enameindividually. An isolating sealaon all contact surfaces at asse

bly and on all fasteners. The ftings and attachments are fillesealed around their peripherieCorrosion-resistant steel or titnium fasteners are used exclusively with carbon components

TITANIUM

Titanium use has greatly in-creased. Titanium alloy forgingare used in the main landing gsupport structure and for vario

fuselage and nacelle strut fittiIn addition, titanium is used fohigh-pressure tubing and ductand for firewalls, door threshoand scuff plates.Large quantities of coated titafasteners (treated to preventcorrosion) are used, includingsome new types especially deoped for use in composite struture.

Composite Material Usage

EAT doc 51 Issue 1



7/167


Wing

The wing surfaces develop aerody-namic forces that support the

airplane in flight. The wing storesfuel, houses the fuel systemequipment, supports the engines,and contains the flaps, spoilers,and ailerons.

Location references on the wingare indicated by distance, ininches, from a base point along aspecific reference line. Wing sta-tions (WS) are measured perpen-dicular along the rear spar. Wingbuttock lines (WBL) are measured

parallel from the fuselage center-line.

WING PRIMARY STRUCTURES

The wing primary structures arealuminum. They are the front andrear spars, upper and lower sparchords, webs, skin panels andstringers, and r ibs. The upper andlower spar chord extrusions attachto

the front and rear spar webs.Chords, stiffeners, and webs makeup the ribs. Conventional ribs arespaced through the entire wing.

Shear tie ribs distribute specificloads to the wing frame. Fuelbaffle ribs minimize fuel slosh inthe fuel tanks. Tank end ribs aresealed and form the ends of thefuel tanks. Side-of-body ribs jointhe outboard wing sections to thecenter wing section. Upper andlower aluminum skin splice platesjo in the sk in panels. Upper andlower aluminum stringersstrengthen the skin panels. Thelanding gear is supported by the

landing gear support beam andrear spar.

WING SECONDARY STRUCTURES

The secondary structures, whichsupport aerodynamic fairings orskins, flight control surfaces, andcontrol mechanisms, consist of theleading edge, trailing edge, and

wingtip. The leading edge is clevered forward from the frontand is made of aluminum ribs skin panels. The leading-edge

slats attach to the leading edgThe trailing edge is cantilevereaft from the rear spar and supports the flaps, aileron, and spers. The wingtip is an aerodynfairing covering the outboard eof the wing. Navigation lightsattach to each wingtip.

The wing, outboard of theside-of-body rib, has access hin the lower surface between rSimilarly, the wing center sect

has a single hole just to the r iof the keel beam and one accopening in each of the threespanwise beams. The dry bay each engine has four accesspanels to the wing tank, and ttwo ribs immediately outboardthe side-of-body splice have acess openings.

Wing Structure

EAT doc 51 Issue 1



8/167


The access openings allow inspec-tion, maintenence, and repair ofinternal wing structure, fuel tanks,and system components.

Trailing edge flight control surfaceshave skin panels made of ad-vanced composites. The spoilerand aileron are carbon-epoxy andtrailing edge flaps are aramid.Structural ribs are made of alu-minium.

Wing Center Section

The wing center section is en-closed within the fuselage an

consists of upper and lower skinpanels and front and rear spars.Other structural members areupper and lower spar chord extru-sions, stiffeners, webs, and floorbeams. Throughout the wing cen-ter section the skin panels arereinforced by spanwise stringersand the spars are reinforced byvertical stiffeners. Spanwisebeams are made of stiffeners andwebs.

Floorbeams are made of chord,stiffeners and web chords.

The wings are attached to thecenter section with the front arear splice fittings, lower side-body splice.. The spar splice

fittings are vertically mountedsections machined from the aminium alloy. The lower side sof-body splice is double-shearsplice. The splice is an aluminchord . The left and right mainsections are joined by more th400 bolts per joint to the wingcenter section to form a unit.

Wing Centre Section

EAT doc 51 Issue 1



9/167


Horizontal Stabilizer

The horizontal stabilizer hassimilar left and right outboard

sections spliced on the airplanecenterline. The main torque boxfront and rear spars have ma-chined aluminum chords and webs.Aluminum ribs join the front andrear spars. Upper and lower alumi-num skin panels are fastened toaluminum stringers attached to theribs. The outboard forward torquebox is between the auxiliary andfront spars. The auxiliary spar hasaluminum extruded chords andclad sheet webs. Aluminum ribs

join the auxi liary and front spars.Upper and lower skin panels arealuminum face sheets over analuminum honeycomb core.

The fixed trailing edge is made ofstiffened ribs covered with skinpanels.Ribs are aluminum alloy, and thepanels are carbon-aramid-fiber-glass hybrid with an aramidhoneycomb core.

An aerodynamic seal extends aftfrom each fixed trailing edge to theelevators. The removable leadingedge is made of aluminum honey-

comb sandwich panels attached tothe auxiliary spar.

Horizontal-stabilizer-to-body seal-ing doors are between the fuse-lage and inboard side of the hori-zontal stabilizer. The sealing doorsare fiberglass panels supported byaluminum alloy ribs.

The horizontal stabilizer is at-tached to fuselage structure bytwo pivot bearings mounted off the

rear spar and the jackscrew at thecenterline of the front spar. Thejackscrew mechanism pivots theentire stabilizer up or down aboutthe two pivot bearings at the rearspar.

The elevators are made fromcarbon epoxy honeycomb panspars, and ribs. Three actuatomove each elevator on eight

hinges.

Access to the center sectiontorque box is through the fronspar closeout panel between erib. Each closeout panel has tinspection holes. Inspection ooutboard main torque box isthrough holes in the rear sparlower surface of the trailing edhas hinged doors and removapanels for access to elevatorhinges, actuators, control link

hydraulic lines, and wire bundBehind the removable leadingare inspection holes in the auxiary spar. The stabilizer tip isremovable to expose the tip r iThe tip rib has inspection holeview the outboard ends of thehorizontal stabilizer.

Horizontal Stabilizer Structure

EAT doc 51 Issue 1



10/167


Vertical Stabilizer

The main structural components ofthe vertical stabilizer are the

forward and main torque boxes,fixed trailing edge, removableleading edge, fin tip, dorsal fin,and rudder.

The forward torque box is betweenthe auxiliary and front spar, andthe main torque box is between thefront and rear spar. Auxiliary, front,and rear spars are aluminum. Thespars have chord extrusions withsheet webs. Aluminum ribs fitbetween the spars. The main

torque box is made from aluminumskin panels riveted to aluminumstringers attached to the spars andribs.The fixed trailing edge has alumi-num ribs covered withcarbon-aramid-fiberglass hybridskin panels. A removable fin tipattaches to the top of the verticalstabilizer. The fin tip is an alumi-

num frame structure with alumi-num and fiberglass-aramid skinpanels. The dorsal fin has alumi-num frames covered with alumi-

num skin panels. An aerodynamicseal closes the gap betweenrudder leading edge and verticalstabilizer trailing edge.

The rudder is hinged to the verticalstabilizer fixed trailing edge ateight places. Three hydraulicactuators move the rudder. Therudder is made of carbon-epoxyhoneycomb sandwich panelsattached to two carbon spars andeight ribs. The tip is

aramid-fiberglass material.The vertical stabilizer is enteredfrom the fuselage through thebody-to-stabilizer access panel.Above the body-to-stabilizer ac-cess panel the ribs have openingsto allow access into the aft torquebox. Access to TV or HF couplers,feedline, and TV antenna isthrough access panels on the

forward torque box. Removingsections of the leading edge gaccess to inspection holes in auxiliary spar. Removing secti

of the fin tip allows access to VOR antenna. Removable panin the upper rear spar allow viing inside the aft torque box.Removable panels access therudder hinges. Rudder controland actuators are accessiblethrough doors in the trailing edleft side. The forward torque binspected through access panand openings from the aft torqbox.

Vertical Stabilizer Structure

EAT doc 51 Issue 1

Revision 00 date 01/06 Pag


11/167

Flight DFlight DFlight DFlight DFlight D

Features

The 757 flight deck featuresstate-of-the-art displays and digitalelectronic systems that allow thetwo-member crew to function assystem managers. The 757 hasone of the most advanced flight

decks ever developed, withsolid-state electronic instruments,cathode ray tube (CRT) displaysfor instant flight information, auto-matic navigation and landingsystems, and improved flight crewvisibility.The latest digital technology withcontrol-display integration providesfor uncluttered instrument panels,optimized crew workload, andimproved operational capabilities.

Displays are designed to shownecessary information but also tomake more information than everavailable to the pilot on command.

757 AND 767 SIMILARITIES

By design, the flight decks for the757 and 767 are so similar thatonly the experienced eye can tellthem apart. Common handlingcharacteristics and display indica-tions, recordings, aural warnings,

and nomenclature are fundamen-tally identical.

Flight decks for both airplanesfeature the same arrangement andlocation of windshields and win-dows, uncluttered instrumentpanels, and just the right balanceof technology.

SIMPLIFIED CAUTION ANDWARNING SYSTEM

Visual, aural, and tactile signalsalert the flight crew to conditionsrequiring their attention. Thesystem was simplified by minimiz-ing the number of different aural

alerts, which are grouped accord-ing to the level of action andawareness required, and by reduc-ing nuisance alerts.

LOW-NOISE WINDOWS

The curved side windows on the757 are designed to reduce aero-dynamic noise and resultingspeech interference in the flightdeck-another contribution from a

carefully planned research anddevelopment effort focused onimproving crew safety and perfor-mance.

CREW SEATING COMFORT

More durable crew seats withcompletely adjustable backs areprovided to further improve crewcomfort.

Design

Panel Arrangement

Lighted Pushbutton Switche

Lighting

Crew Seats

Windows

Crew Oxygen system

EAT doc 51 Issue 1



12/167


Design

The design of the 757/767 flightdeck is the result of a long and

carefully planned program ofresearch and development. Theobjective was to provide a flightdeck that would meet the needs ofairline flight crews through the1990s and beyond. The goals ofthat program were:

Safety.

Improved operational capa bilities.

Optimized performance.

Reduced workload. Reliability.

Maintainability.

Low operation costs.

The technology used to meetthese goals included:

Digital computers and micropro-

cessors.Color cathode ray tube (CRT)displays.Integrated flight managementsystems.Laser gyro inertial referencesystem.Advanced systems monitoring.Built-in test equipment (BITE).

The Boeing flight deck designphilosophy provides enhancedsafety and productivity through

improved crew comfort and perfor-mance and optimized workload.

Crew comfort is improved byproviding more comfortable andurable seats, lower noise levmore efficient air-conditioning

better internal and external vis

Crew performance is improveda result of modern-technologyelectronic flight instruments foorientation, a flight managemesystem for airplane navigationperformance optimization, a ctralized crew alerting system, uncluttered instrument panels

An optimum crew workload levneither so high as to cause ov

work, nor so low as to causelapses in attention. System deuses simplification, automatioand redundancy to provide asimple man-machine interfacewhile maintaining the proper cemphasis on flightpath contro

Flight Deck Panels

EAT doc 51 Issue 1



13/167


The flight deck is designed to bequiet and dark, with safety andproductivity increased through acaution and warning system,

improved crew comfort, and re-duced workload. The caution andwarning system reduces nuisancealerts and the number of differentaural alerts. Aural alerts are cat-egorized according to the level ofcrew action and awareness re-quired, and no immediate crewaction is required after the firstfailure within a subsystem. Crewcomfort and reduced workload areattained by providing adjustable

seats, lower noise levels, moreefficient air-conditioning, bettervisibility, simplified procedures,accessibility of all controls toeither pilot, simplified systemdesign, and elimination of itemsused for maintenance only.

Panel Arrangement

The 757 flight deck has a commonflight crew type rating with the 767.

The 757 and 767 share similarhandling characteristics, check-lists, and visual aler ts, and havethe same crew procedures; recallitems; aural warnings; flight deckarrangements; windshield; panellocation, arrangement, and nomen-clature; and controls. Initial andrecurring training will qualify crewfor both airplanes.

The 757-200 Freighter flight deck

includes a double crew seat, aflushing toilet, and a rigid cargobarrier.

MainInstrument Panel Arrangement

EAT doc 51 Issue 1



14/167


Lighting Control Panel

Captains and First Officers Panel

Glearshield Panel

Main Panel Center

Main Instrument Panel

EAT doc 51 Issue 1



15/167


Overhaed Panel

1. Inertial reference system (IRS) control and display2. Yaw damper3. Hydraulics

4. Miscellaneous alert annun ciator lights5. Stand-by power6. Electrical7. Auxiliary power unit (APU) ignition

8. Cockpit recorder 9. Lighting control panels10. Emergency lights, passen ger oxygen

11. Ram air turbine, engine start12. Fuel system13. Fuel quantity14. Anti-ice15. Windshield wiper

16. Window heath17. Selective calling (SELC18. Passenger sign19. Cabin altitude controls

20. Cabin pressure gauge21. Equipment cooling22. Compartment tempera23. Zone temperature/pac control24. Bleed air system

KEY:

EAT doc 51 Issue 1



16/167


Control Stand

EAT doc 51 Issue 1



17/167


Right Side Panel

Basic

Optional

EAT doc 51 Issue 1



18/167


Lighted Pushbutton Switches

Alternate action push-buttonswitches and lights provide control

inputs and status indications. Bothalternate action and momentarypush-button switches and lightsare used on the 757.

ALTERNATE ACTION SWITCH

All alternate action switches aremechanically latched to the lastoperated position (in or out). Eachsucceeding operation selects theswitch to the opposite position.The switch position is indicated bythe absence or presence of amechanical flag in the switch face.The switch position display (flag)has a white legend on a blackbackground in the latched INposition and is illuminated by5V-ac, 400-Hz power. The legendis hidden by a mechanical shutterin the OUT position.

MOMENTARY SWITCH

Pressing the momentary switctransfers the switch contacts.

switch does not have switch ption display; however, the lightdisplay can indicate the positia relay or contactor controlledthe switch.

STATUS/CAUTION DISPLAYS

The status/caution display porof the switch is a light that disa system condition. A legend cbe either a color or black on ea black or white background.Indication lights use the mastedim and test system power,26.5V-do bright and 12V-do d

Lighted Pushbutton Switch (Mechanical)

Alternate action type Momentary type

EAT doc 51 Issue 1



19/167


Lighting

General Illumination of the flightdeck is provided by ceiling-

mounted dome lights. The domelights are controlled by a rotarydimmer switch on the overheadpanel (P5). Specific area illumina-tion is provided and controlled ateach flight crew station by map,chart, and portable utility lightsPanel illumination is provided bypanel lightplates that are con-trolled by a rotary dimmer, Theindicator lights incorporate a dimand test feature. An override

switch is provided to illuminate thedome and floodlights in the britmode.Circuit brakers, relays, and dim-ming cards for the flight decklighting are located in the lightingpanel (P26)

Flight Deck Ligting

EAT doc 51 Issue 1



20/167


Crew Seats

The captains and first officersseats move on curved tracks to

facilitate ingress and egress.Armrests, seat backs, and thighpads are manually adjustable forvariations in personnel size. Seatshave lap belts, crotch straps, andshoulder harnesses attached toinertia reels. The first and secondobservers seats are bulkheadmounted and are not adjustable,and fold up when not in use.

Stowage space for suitcases iunder the observers seat. A ccloset is located to the right aaft of the first officers seat. F

kits stow outboard of the pilotsseats.

Miscellaneous equipment andfurnishings in the flight deckinclude crew equipment consoglareshield and sunvisors, ashtrays, smoke goggle stowagepockets, hand microphones, hsets, oxygen masks, observerpanels, removable wastebaskcupholders, and a spare light holder.

Flight Deck Arranngement

EAT doc 51 Issue 1



21/167


Windows

Flight deck windows consist of twofixed, flat forward windowshields;

two sliding, curved side windows;and two fixed, curved side win-dows. The two sliding windows(No. 2 windows) serve as emer-gency exits and are replaceablefrom the inside. The No. 1 flatwindshield and the No. 3 fixed,curved windows are replaceablefrom the outside. The windows areheated electrically to providedefogging and anti-icing.

The curved windows significantly

reduce the aerodynamic noisecontribution to speech interferencelevel (SIL) in the flight deck.

The eye position indicator attachesto the windshield center post andhas two sighting points. Thisenables the pilot to adjust the seatto the most advantageous positionfor viewing instruments and theoutside.

Flight Deck Windows and Noise Control

EAT doc 51 Issue 1



22/167


Crew Oxygen System

A diluter demand oxygen systemsupplies the flight crew with oxy-

gen and includes an oxygen maskwith an integral pneumatic harnessat each crew-station. The maskassembly includes system test andselection functions.

Gaseous oxygen is supplied from ahigh-pressure cylinder located inthe forward lower cargo compart-ment immediately aft of the cargodoor.The high-pressure oxygen isreduced by a pressure regulator

and supplied to the flight deck. Thecylinder contains a shut-off valve,thermal relief, and a pressuregauge. Attached to the cylinder isa pressure regulator and a pres-sure transducer. The pressuretransducer provides a signal to theEICAS, which provides a display ofcylinder pressure on the statuspage. The oxygen is supplied toeach crew-station, which containsa diluter demand mask regulator.

The mask regulator is stored icontainer and can be tested wout removal of the mask.

Overpressure in the oxygen bcauses the thermal relief diskrupture, discharging the conteof the bottle overboard. A therrelief indicator on the right sidthe fuselage shows that this hoccurred.

Crew Oxygen System

EAT doc 51 Issue 1



23/167

Electrical PElectrical PElectrical PElectrical PElectrical Powowowowow

Features

AC POWER

Primary electrical power for the757 is generated by the integrateddrive generator (IDG) mounted oneach engine or from the generator

driven by the auxiliary power unit.

Each generator produces 90 kVA,115/200V, 3-phase, 400-Hz powercontinuously and is capable ofcarrying all essential loads.

DC POWER

If primary power fails, essentialloads automatically transfer to the

backup power 28V DC battery and115 V, single-phase, 400-Hz staticinverter. Normal 28V-do power issupplied through the transformer/rectifier unit.

AC Power Overview

DC Power Overview

System Control and indication,

Hydraulic Motor Generator

Electrical System Panels


The 757 and 767 electrical powersystems were certified as essen-tially the same. All major equip-ment is identical or similar, as inthe case of the IDGs, which haveonly minor engine-dependent

differences.

Equipment failure rates and prob-ability of loss of power sources areessentially the same for bothairplanes.

The bus configuration of the acpower system, main do system,standby power system, and hy-draulic motor generator system (ifinstalled) are also essentially the

same for the two airplanes.

EAT doc 51 Issue 1



24/167


The 757 electrical power systemsare designed to supply airplaneuser systems with alternatingcurrent (ac) and direct current (dc)

power.

AC Power Overview

The ac power for airplane groundoperations is supplied through theexternal power panel or from theauxiliary power unit (APU)-drivengenerator. For in-flight operations,power is supplied from an inte-grated drive generator (IDG)mounted on each engine or fromthe APU-driven generator. Each

generator can supply 90 kVA, 115/200V, 400 Hz, 3-phase ac powerand cannot be paralleled. Majorcomponents associated with the acsystem include three generatorcontrol units (GCU), a bus powercontrol unit (BPCU), and powerpanels located in the main equip-ment center. An optional electricalgenerating system, the hydraulicmotor generator (HMG), operatesas a non-time-limited backup

source in the event of loss of allmain electrical power.

DC Power Overview

Normal airplane 28V-do power isproduced by AC/DC conversion.Battery systems provide alternatedo and standby power. Major dosystem components include a mainbattery, battery charger, twotransformer/ rectifier units (TRU),and static inverter. Componentsused with the APU do system-APUbattery, charger, and TRU-arelocated in the aft equipment cen-ter.

System Control and Indication

The electrical system controlpanels provide manual or auto-matic source selection. A momen-tary test switch is provided forHMG system checkout. Electricalsystem monitoring is provided byEICAS messages.

AC POWER

The main ac buses supply all the essential ac loads in the a

plane. Each bus is divided intoindependent sections. An ac tbus provides interconnectionbetween the main buses undecertain conditions. The utilitybuses supply nonessential loasuch as passenger entertainmand reading lights. Galley powalso considered a nonessentiaload. Nonessential loads can sautomatically to protect powersources. The optional APU TRenergizes the APU starter mo

the right main ac bus is energduring APU star ting. The APUbattery starts the APU if an ATRU is not installed.

Electrical System Overview

EAT doc 51 Issue 1



25/167


The ground service bus suppliesboth in-flight and ground loads,including interior lights, batterychargers, and cooling fans. The

ground handling bus suppliesloads that are used only duringground operations, such as cargohandling equipment. This bus isonly powered on the ground.

The center buses supply both acand do power to the center chan-nel equipment of the autolandsystem. During Cat III autolandoperation, the buses are suppliedfrom alternate sources indepen-dent of the main buses.

The flight instrument transferbuses supply power to selectedcaptains and first officers flightinstruments and allow automatictransfer to an alternate powersource in case of normal sourcefailure.

The ac standby bus supplies singlephase power to essential flightloads and automatically transferspower sources in case of primary

source loss.

DC POWER

The left and right do buses supplypower to loads requiring do power.Each main do bus is divided intoindependent sections. When eitherbus is unpowered, the do tiecontrol unit automatically ener-gizes the do tie relay.

The DC standby bus supplies

power to certain essential airplaneloads and transfers sources incase of primary source loss.

The DC ground handling bussupplies do power for groundhandling equipment and is ener-gized on the ground only.

A main battery and battery chsystem provides a dedicatedsource of power for operation the standby and autoland syst

The separate APU battery andbattery charger system providpower for APU starting.

Hydraulic Motor Generator(Option)

For extended-range twin operations (ETOPS), the HMG systprovides a non-time-limited alnate source of ac and do powafter loss of all generator powflight. An ac generator supplie

captains flight instrument andand right transfer buses. A recfied generator output powers thot battery bus.

Electrical Power

EAT doc 51 Issue 1



26/167



Normal operation of the electricalpower system is performed at the

electrical system control panel.Both momentary and alternateaction push-button switches areused on the electrical panels. Thealternate action switch is mechani-cally latched to the last operationposition (in or out). Switch positionis indicated by the absence orpresence of a mechanical flag inthe switch face. Indicator lights arepowered by the master dim andtest system.

ELECTRICAL SYSTEM CONTROLPANEL (P5)

The momentary external powerswitch controls opening and clos-ing of the external power contac-tor. A white AVAIL light indicatespower is of proper quality. Thewhite ON light illuminates when-ever the external power contactoris closed.

Generator control switches providea control signal that closes thegenerator control relay (GCR) and,when proper power is available,

closes the generator circuitbreaker (GCB). The flow bar andON legend indicate switch position.An amber OFF light illuminateswhen the associated generatorcircuit breaker is open.The ac bus tie switches allowmanual or automatic control of thebus tie breaker (BTB). In theunlatched position the associatedBTB is opened, isolating theassociated main ac bus from theac tie bus. The AUTO indication is

not visible and the amber ISLNlight is illuminated. Operating theswitch to the latched-in position(normally AUTO illuminates andISLN extinguishes) enables auto-matic operation of the bus tiebreaker. If ISLN illuminates whenthe bus tie switch is latched, afault has tripped and locked theBTB open.

The AC BUS OFF indicator ligilluminates when the main ac is de-energized.The latching utility bus switch

provide manual control of thepower relays connecting utilitygalley buses to the left and rigmain ac buses. The ON legendindicates switch position and ihidden when the switch is in thout position. The amber OFF lilluminates if the associated ubus relay is open.

The momentary generator drivdisconnect switches cause amechanical disconnection betw

the IDG and the engine. Theswitches are spring loaded to out position and illuminate witoil pressure or high oil temperture.


EAT doc 51 Issue 1



27/167


The battery switch controls con-nection of the battery bus to theleft do bus or the hot battery bus.The ON legend indicates switch

position and is hidden when theswitch is in the out position. AmberOFF light illuminates when thebattery switch is in the out positionduring normal ground and flightoperations.A battery discharge light illumi-nates if the battery is discharging.The standby power selector switchcontrols standby power mode. Thestandby system is turned off bypushing the switch in and turning itfrom the AUTO to the OFF posi-

tion. An amber standby power busoff light illuminates when the ac ordo standby bus is unpowered.The standby power selector switchcan be turned to the BAT positionto test the output of the mainbattery and static inverter.

AUXILIARY ELECTRICAL SYSTEMCONTROL PANEL (P61)

The momentary generator field

manual reset switch opens orcloses the generator (field) controlrelay (GCR) if the generator con-trol switch is unlatched. A whiteFIELD OFF light illuminates whenthe GCR is open.

Built-In Test Equipment (BITE) Display

EAT doc 51 Issue 1



28/167

AuxiliarAuxiliarAuxiliarAuxiliarAuxiliary Py Py Py Py Power Systower Systower Systower Systower Syst

Features

OPERABLE ON THE GROUND ORIN FLIGHT

The auxiliary power unit (APU)provides an alternate powersource to support aircraft systems

on the ground or in flight. It alsoprovides pneumatic power forenvironmental control system andmain engine star t. All APU opera-tions are governed and coordi-nated by the electronic control unit(ECU), which includes extensivebuilt-in test equipment for faultdiagnosis and protective shut-down.

AIRPLANE SELF-SUFFICIENCY

The APU supplies pressurized airfor engine starting and for main-taining cabin air-conditioningduring ground operations.

AUTOMATIC SHUTDOWN FEATURE

The APU shuts down automaticallyto prevent damage fromoverspeed, high oil temperature,low oil pressure, or a blockedgenerator oil filter.

OPERABLE DURING REFUELING

The APU allows air-conditioning orelectrical power to be suppliedduring refueling.


The 757 and 767 use exactly the

same APU. An hour meter is basicon the 757 and optional on the767. Basic equipment also in-cludes automatic low oil quantitydiscrete light and message onEICAS.

The few differences between thetwo models include circuit breakernomenclature, drain tube differ-ences with different drain mastlocations, air intake system design

and materials, exhaust duct as-sembly, and thermal insulationblanket. Ground signature pins inthe APU controller provide fordifferent reverse flow shutdown onthe 757 and higher bleed air capa-bility on the 767.

Auxil iary Power system

Ind ication

Lubricat ion System

Fue l Sys tem

Ignit ion Starting System

Pneumat ic System

Electronic Control Unit Inpu

Output

EAT doc 51 Issue 1



29/167


Auxiliary Power System

The auxiliary power system sup-plies electrical and pneumatic

power for the airplane. On theground, electrical and pneumaticpower makes the airplane indepen-dent of ground support equipment.

The auxiliary power unit (APU) is aGarrett GTCP 331-200 controlledby an electronic control unit (ECU).The ECU is located in the E6 rackwith the APU battery, batterycharger, and the optional APUstart transformer/ rectifier unit(TRU).

The ECU coordinates the startingsequence, monitors the operationand pneumatic output of the APU,and ensures proper shutdown. TheECU features extensive built-intest equipment (BITE) that moni-tors many line-replaceable units(LRU) and initiates protectiveshutdowns to prevent damage tothe APU.

These shutdowns and failed com-ponents are identified on the faceof the ECU.

The auxiliary power system iscontrolled from the APU controlpanel on the P5 panel. This panelfeatures a three-position rotaryswitch and fault and run annuncia-tor lights. The engine indicationand crew alerting system (EICAS)shows APU exhaust gas tempera-ture (EGT), revolutions per minute(rpm) in percent speed, and oilstatus. An APU hour meter andoptional cycle meter are locatedon the P49 panel. To shut down the

APU normally, the control switch isturned to OFF. To shut down theAPU during an emergency, theAPU fire handle on the P8 panelmust be pulled or the APU shut-down switch on the APU remotecontrol panel (P62) located on theright side of the nose gear acti-vated. When the APU is shut downusing the P62 APU shutdownswitch, the battery switch in the

flight deck must be cycled off on before the APU can be sta

The APU is warranted to start

to an altitude of 35,000 feet. Icapable of supplying 115V-ac3-phase electrical power up toservice ceiling of the airplane.Pneumatics is available up to altitude of 17,500 feet. If bothelectrical and pneumatic demaare present, the ECU reducespneumatic output as necessarprevent exceeding APU EGT lThe ECU senses five differentpneumatic modes of operationfrom the airplane pneumatic s

tems.

The ECU positions the inlet guvanes (IGV) in response to themodes to ensure efficient opetion and load compressor surgcontrol.

Auxiliary Power System

EAT doc 51 Issue 1



30/167


APU Systems and Components

APU Installation

EAT doc 51 Issue 1



31/167


Indication

The ECU sends analog signals tothe EICAS computers for percent

rpm and EGT and discrete signalsfor fault shutdowns and for somefaulty LRUs stored in nonvolatilememory (NVM). In addition, the oilquantity transmitter sends ananalog signal of oil level directly tothe EICAS.

APU speed in percent rpm anEGT in C are displayed on thEICAS STATUS and PERF/APpages.

The advisory message APU FAappears and the FAULT lightilluminates to annunciate a protive shutdown of the APU. Thefault light is also used to showtransit of the APU fuel shutoffvalve.

The white RUN light illuminatethe APU control panel whenevthe APU is operating above 95speed.

An hour meter and optional cymeter are located on the P49panel in the aft equipment cento record operating hours andcycles, respectively. The APUbattery bus powers the hour mWhen APU speed is greater th95%, the ECU provides a grouto operate the hour meter.

APU Indication

EAT doc 51 Issue 1



32/167


Lubrication System

The APU lubrication system con-sists of an oil supply; a pressure

system for oiling the bearings,generator, and star ter clutch; ascavenge system for returning oilto the sump from the bearings; agenerator oil scavenge system; agearbox pressurization system;and an oil cooler with bypass.

The APU gearbox serves as an oilreservoir. Servicing is by apour-type fill port. Oil quantity isindicated by a sight glass and anoil quantity signal to EICAS. Mag-

netic chip detectors are alsoinstalled.

A gear-type oil pump in the gear-box sends pressurized oil throughan oil cooler and filter to thebearings and generator. When theoil is cold, a deoil solenoid valveopens, allowing the pump to drawair from the gearbox, which de-creases the oil drag and enableseasier starting. Low oil pressure

switch and oil temperature sensorsignal the ECU, causing protectiveshutdowns if limits are exceeded.

Oil cooler is located between theoil pressure pump and bearings.An oil cooler bypass valve sendscold oil around the oil cooler. Thisvalve also allows bypass of anobstructed cooler. Two checkvalves prevent backflow and draindown.

Three scavenge pumps return oilto the reservoir. The compressorbearing scavenge pump and gen-erator scavenge pump are

positive-displacement gear type.The turbine-bearing scavengepump is a gerotor type. Scavengeoil from the generator flowsthrough a non-bypass filter toprotect the APU from oil contami-nation if the generator fails. Agenerator oil filter differentialpressure switch signals the ECU ifthe generator oil filter becomesobstructed. This initiates a protec-tive shutdown.

At higher altitudes (around 18feet), the low ambient air prescould cause oil foaming. Thegearbox pressurization system

prevents this by pressurizing tgearbox with second stage copressor air; Pcd2 Componentsinclude a gearbox shutoff valvshuttle valve, and a gearbox psure-regulating valve. Operatioautomatic and controlled pneumatically.

Protective shutdowns that areassociated with the lubricationsystem are low oil pressure (Lhigh oil temperature (HOT), an

blocked generator oil filter (GEFILTER).

The faulty units stored in the Ememory with respect to the lucation system are LOP SWITCDEOIL SOL, HOT SENSOR, aFILTER SWITCH.

APU Lubrication System

EAT doc 51 Issue 1



33/167


Fuel System

The APU fuel system receives fuelfrom the airplane wing tanks

through a shrouded line. Thesystem then pressurizes, filters,and meters fuel for combustionand to operate the inlet guide vaneactuator (IGVA). The primarycomponents are the fuel controlunit, flow divider, primary andsecondary fuel manifolds andnozzles, and IGVA. The APU is aconstant-speed engine. Speedcontrol is accomplished automati-cally by the ECU through torquemotor inputs to the fuel control

unit, resulting in fuel flow regula-tion. The acceleration schedule isalso torque motor controlled.

Air inlet pressure (P2) and inlettemperature (T2), or load com-pressor inlet temperature (LCIT))are sensed by the ECU to adjust

fuel flow for ambient conditions.The torque motor also responds toTS (EGT limits), if necessary, toprevent an OVERTEMP protectiveshutdown.The fuel control unit accomplishesall pressurizing, filtering, andmetering for the APU. It mounts tothe front of the oil pump. Electricalconnections include the torquemotor and fuel shutoff valve sole-noid, which are ECU controlled.

The fuel flow divider separates themetered flow into two manifolds:primary and secondary. The pri-mary manifold is used full time.The secondary manifold is usedwhen flow demands are increased.An ECU-controlled electric sole-noid valve modifies secondaryflows to accommodate APU start-ing requirements.

The two fuel manifolds encirclAPU combustion chamber. Eahas six fuel nozzles permanenattached. The nozzles and ma

fold are an LRU as an assembonly.

APU protective shutdowns thaassociated with the APU fuelsystem include NO FLAME, NACCEL, SLOW START,OVERTEMP, and OVERSPD. NACCEL and SLOW START areoften caused by too little fuel,while OVERTEMP and OVERSare often caused by excessiveflows.

Faulty LRUs that can be displon the face of the ECU with respect to the fuel system are FCONTROL, FUEL SOL, and FDIV SOL.

APU Fuel System

EAT doc 51 Issue 1



34/167


Ignition/Starting System

The ignition/starting system pro-vides initial APU acceleration and

combustion spark. The systemconsists of the ignition unit, theigniter, and the star ter motor.

The ignition unit provides igniterspark energy. The igniter providesthe spark to the combustion cham-ber. Ignition unit power is con-trolled by the ECU.

The starter motor provides APUinitial rotation and acceleration. Itis powered by the APU battery or

the opt ional APU TRU. AC powersense relays determine the powersource used.

The main battery switch must beon to start the APU. APU start isinitiated by rotating the APU startswitch momentarily to START andreleasing it to ON.

Start initiation opens the APU airinlet door. Once the door is open,the ECU energizes the APU crankcontactor or optional TRU start

relay, as appropriate, to supplypower to the star ter motor. At 7%speed, the ECU energizes theignition unit. At 50% speed, theECU de-energizes the startermotor. At 95% speed, the ECUde-energizes the ignition unit.

APU Ignitio/Starting System

EAT doc 51 Issue 1



35/167


Pneumatic System

The APU is designed to providepneumatic power to the airplane

for environmental control system(ECS) and main engine start(MES). A gearbox-driven fan blowsair through the oil cooler and intothe APU compartment for thecooling system. All air is drawnfrom the inlet door and ducting intothe plenum for these systems.

Plenum air is drawn through vari-able IGVs to the load compressorand is discharged into the airplanepneumatic ducts. The IGV s are

essentially a pneumatic valve,designed to control the volume ofair available to the compressor.This improves the efficiency of theAPU by unloading the load com-pressor when airplane pneumaticsare not demanded.

The IGVs are controlled by theECU through a torque motor in theIGV actuator. Fuel pressure isused for actuation power. Feed-

back is through a linear variabledifferential transformer in theactuator.

The load compressor output air-flow must match the input orsurges may occur, causing erraticand damaging operations. A modu-lating surge control valve is de-signed to dump excess air (notrequired for airplane pneumatics)to prevent this surging. The valveis modulated by the ECU as a

function of air mass flow in theoutput ducting, and a delta-P flowsensor system is used to signalthe ECU for this pur pose. The flowsensor consists of a static pres-sure ring (PS), a total pressuresensor (PT), transducers, and avariable volume chamber. Valvemodulation is by torque motorcontrol and Pcd2power.

The ECU controls the IGVs athe surge valve as a function airplane pneumatic demand msignals from the ECS, sensor

signals, and the settings onswitches located behind a platthe face of the ECU. Theseswitches allow adjustments tosoftware without internal reprogramming to accommodate vaable external operational circustances. The switches are notdesigned for calibration or onladjustments.

Air from the plenum is drawn bthe gearbox-driven fan throug

fan isolation valve to cool the and the APU compartment. Thfan isolation valve is pneumatopened during APU start whePcd2 reaches 7.5 psig. Air theblows continuously through thcooler and compar tment. A thmal bypass valve prevents colfrom flowing through the coole

APU Pneumatic System

EAT doc 51 Issue 1



36/167


A REVERSE FLOW protectiveshutdown occurs if the compressorsurges. This also protects the APUfrom upstream pneumatic system

failures (check valves and bleedvalves) that would allow mainengine air to flow back through theload compressor. The INLETDOOR protective shutdown en-sures that the APU has sufficientincoming air by not allowing a startuntil the inlet door is open. TheLCIT SENSOR, ECS CONTROL,IGV ACT, FAN VALVE, PT SEN-SOR, DELTA-P SENSOR, andSURGE VALVE are identified asfaulty units when necessary.

Electronic Control Unit Inputand Output

The ECU may be powered by

turning the APU control switch toSTART or, when this switch is off,by activating one of the threetoggle switches on the face of thecontroller. The controller automati-cally powers down when the APUcontrol switch is off, APU rpm isbelow 7%, and BITE proceduresare complete.The ECU receives analog anddiscrete input from the airplaneand the APU.ECU output includes EGT and rpm

signals to EICAS, aircraft discretesignals, and APU signals, bothanalog and discrete, for torquemotors and solenoids.Normal operation of the APU andECU is completely automatic whenSTART has been selected on theAPU control switch. Once the APUis on speed (over 95% rpm), theoperator may elect to draw electri-cal power and pneumatics asdesired. The controller automati-

cally performs system monitorand protective shutdown functThe requirement to interrogateECU for fault information is an

ciated in the flight deck by theAPU FAULT ligh t, by the APUFAULT EICAS advisory messaand by the APU BITE EICASmessage on the ECS/MSG paSome fault LRUs do not inhibAPU operation or cause a protive shutdown. The APU BITEEICAS message appears for o12 of the 24 LRU faults. In gethose faulty LRUs that manifethemselves by other indicationsuch as through a protective

shutdown or loss of pneumaticoutput, do not cause the messThe instructions for accomplisthe BITE check are on a placaon the door of the E6 rack.

Electronic Control Unit Input and Outpu

EAT doc 51 Issue 1



37/167


The ECU front panel allows faultrecall, reset, and display by use oftwo light arrays and threeswitches. The five-position rotary

FAULT SELECT switch allows theselection of the reason for any ofthe last five protective shutdowns.A toggle switch under the ERASEMEMORY plate clears the faultmemory.

The ECU is not powered unlessthe APU control switch is on or iscommanded by an ECU switch.Prerequisites for BITE interrogation are:

APU or main battery power.

APU control switch in the flight deck off.

APU rpm below 7%.

LAMP TEST. Each column oflamps illuminates in a left-to-rightsequence. If a lamp does notilluminate, the interrogation is not

inhibited, but the fault is not dis-played if present.

FAULT SELECT-FAULT DISPLAY.Place the rotary FAULT SELECTswitch in position 1 and activatethe FAULT DISPLAY switch (up).The most recent protective shut-down illuminates, followed by anyfaulty unit lamp that caused theshutdown. If no protective shut-down is stored, the TST OK lampilluminates. Repeat this procedure

for fault select switch positions 2,3, 4, and 5 to recall the reason forprevious protective shutdowns. TheFAULT SELECT switch operatesonly in conjunction with the FAULTDISPLAY switch.

FAULTY UNIT. Toggle the FAUUNIT switch (down) to annuncall faulty units stored in theFAULTY UNIT lamp array. The

lamps illuminate from top to btom, left to right across the arFaults are not sequenced in thorder sor ted. TST OK illuminano faults are stored.

ERASE MEMORY. Clear all fastored in the ECU by pushing on the toggle switch located uthe ERASE MEMORY cover. TWAIT lamp illuminates while tprocedure is in progress, folloby TST OK.

SELF-TEST. This test is identto the pre-start BITE accomplwhen the APU switch is turnedSTART. WAIT illuminates, folloby any faulty units discoveredsame as in FAULTY UNIT. If nfaults are detected, TST OK ilnates.

Current BTCP 331-200 APU Specifications

EAT doc 51 Issue 1



38/167


39/167


40/167


41/167


42/167


43/167


44/167


45/167


46/167


47/167


48/167


49/167

HydrauHydrauHydrauHydrauHydrau

Features

TRIPLE REDUNDANCY

Three functionally independent,fulltime, 3, 000-psi systems pro-vide hydraulic power for fully

powered flight controls, landinggear, thrust reversers (Pratt &Whitney engines), high-lift, andbraking systems. Hydraulic systemreservoirs are pressurized withbleed air from either engine, theauxiliary power system, or groundair carts.

Distribution systems for the hy-draulics are routed to maximizesystem physical separation.


Hydraulic systems for the 757 and767 are basically identical. Bothhave three independent hydraulicsystems with similar components;only the size differs. The hydraulicsystems for both airplanes aredesigned to operate in the samemanner with full redundancy.

The 757 and 767 use identical

engine driven and electric-motordriven pumps to generate hydraulicpower, and similar ram air turbinesprovide backup hydraulic power tothe center system for primary flightcontrol actuation on each airplane

Hydraulic system servicing is verysimilar because both models haveparts in common, including fillservice, selector valves andground connections. Distributionsystem components such asfittings check valves, and tubingmaterials are identical for nearly

all installations. Titanium tubing isused for pressurized lines. Thefiltration philosophy for both air-planes is similar; with pressureand case drain filters for eachpump and return filters for eachsystem.

Hydraulic system flight deck indi-cations and controls for bothairplanes are nearly identical.Minor differences reflect the two

distinct hydraulic power systemarchitectures, which are based ondifferences in 757 and 767 controlsurface requirements. For ex-ample, the 757 uses only outboardailerons, whereas the 767 usesinboard and outboard ailerons.

The landing gear and high-liftdevices are hydraulically poweredby the left hydraulic system on the757 and the center hydraulicsystem on the 767. An engine

driven pump, an electric motorpump, and a power transfer unitdriven from the right hydraulicsystem if power is lost on the lefthydraulic system supplies powerfor the left hydraulic system on the757. Two electric pumps and anair-driven pump provide power forthe 767 center hydraulic system.

Hydraulic Power Distribution

System Components

Ram Air Turbine

Controls. Indicators. and

Cautions

EAT doc 51 Issue 1



50/167



The three hydraulic systems-left,right, and center-are powered by a

total of seven pumps. Multiplepumps in each system ensurereliability.

There are two levels of redun-dancy. Primary flight controls havethree separate systems supplyingpower-to-power control actuators(PCA) for the control surfaces andautopilot servos. Dual power isused for the elevator feel unit,stabilizer trim, yaw damper servos,and brakes. Thrust reverser, land-

ing gear, and lift device systemsuse a single hydraulic powersource.

The left and right systems aresimilar, with each containing oneengine driven pump (EDP) and onealternating current motor pump(ACMP). A power transfer unit(PTU) connects the left and rightsystems mechanically. A hydraulicmotor in the right system powers a

hydraulic pump in the left systemto provide sufficient flow to retractthe landing gear and operate theflaps and slats in the event of loss

of the left engine or left EDP. Theram air turbine (RAT) retractactuator is powered by the rightsystem. For extended-range twinoperations (ETOPS), an optionalelectric generator driven by ahydraulic motor is required tooperate essential electrical equip-ment in case of loss of electricalpower on both alternating current(ac) buses. The hydraulic motorgenerator is located in the leftsystem and can also be driven by

the PTU.

The center system has twoACMPs for primary pumps and theRAT for emergency power. Thecomponents of the system arelocated in the wheel wells andbody fairings.

There is no fluid interconnectionbetween the three systems.The three independent, full-time,

3000 psi systems use a synthtype IV fluid (BMS 3-11).

Two hydraulic pumps that are

driven from independent powesources normally power eachsystem. Distribution of pressufrom the three systems is sucthat the failure of one system not result in loss of any flightcontrol functions, and the airpcan be safely operated in the of loss of two hydraulic systemAn emergency hydraulic pump(RAT) provides flight control otion in the event of dual enginfailure.

A central fill point facilitates flservicing of all three systems.Reservoir pressurization is obtained from the airplane pneusystem and is available whenethe pneumatic ducts are pressized. External hydraulic powerbe connected to each system.


EAT doc 51 Issue 1



51/167


The three systems are color-codedfor easy identification of tubingand components. The left systemis coded red, the center system

blue, and the right system green.

The components powered by theleft system include flight controls,landing gear, brakes, left enginethrust reverser, hydraulic motor/generator, and nose wheel steer-ing. The left system can be pow-ered by the right hydraulic systemthrough the PTU using reservoirreserve fluid for emergency opera-tion of the landing gear, lift de-vices, and nose wheel steering.

The right system is similar to theleft system. Components poweredby the right system include flightcontrols, brakes, PTU, right enginethrust reverser, and the RAT re-tract actuator. Isolation valves canprovide ACMP output to the brakesonly, using reservoir reserve fluid.

Hydraulic System Schematic

The center system, powered bytwo ACMPs, is smaller than theleft and right systems and powersonly flight controls. The RAT pow-

ers the center system to providehydraulic power for emergencyoperation of the flight controls.Reservoir reserve fluid is for theRAT.

Each reservoir is pressurized from thepneumatic system, and fluid is servicedthrough a common selector valve. Heatexchangers in the main fuel tanks coolcase drain fluid returning to the reservoir.A shutoff valve in each system is used toshut off hydraulic components in the tail

for ground maintenance only.

EAT doc 51 Issue 1



52/167


System Components

The EDPs are variable displace-ment pumps, driven by the engine,

with the output pressure compen-sated to a nominal 3,000 psi. Attakeoff power settings, the pumpcan deliver approximately 37gallons/minute (140 liters/minute).A flight deck ON-OFF switchoperates a solenoid depressuriza-tion valve on the system outputside of the pump. With the switchoff, the pump is depressurized, butfluid flow is maintained through thecase drain circuit for cooling andlubrication. A fire shutoff valve is

located in the fluid supply line tothe pump and closes when theengine fire switch is pulled.

The ACMPs are variable displace-ment, constant- horsepowerpumps that are driven by an elec-tric motor with the output pressurecompensated to a nominal 3000psi.

The pump can deliver approximately 6.7 gallons/minute (25liters/minute) at 2,850 psi andgallons/minute (35 liters/minu

2,000 psi. On-off control of eaunit is provided in the flight deand the pumps are on continuduring normal operations. Whethe pump is operating, positivecase drain flow is maintained cooling and lubrication of thehydraulic pump and electric m

The PTU is a fixed-displacempump driven by afixed-displacement hydraulicmotor. The pump delivers up t

gallons/minute (83 liters/ minuat 2200 psi pressure. The hydpump automatically powers thlanding gear and flap/slats acttion subsystems if the left sysis not operating and the right is operating.

Hydraulic System Component Location

EAT doc 51 Issue 1



53/167


Ram Air Turbine

The RAT is installed in the right aftwing-to body fairing to automati-

cally provide emergency hydraulicpower to the center system flightcontrols in the event both enginesbecome inoperable (rpm below50% in flight). An override switchis provided on the pilots overheadpanel for manual deployment atthe pilotsdiscretion. The turbineand hydraulic pump are mountedon a strut housing that pivots onairplane structure. The RAT com-partment door is opened andclosed by a door actuation link as

the RAT is deployed and stowed.

The actuator is extended byspring force to deploy the RAT andretracted by right hydraulic systempressure to stow the RAT. Retrac-tion can only be accomplished onthe ground. When the RAT isextended in flight, airflow drivesthe turbine, which drives thehydraulic pump.

The RAT is a variable displace-ment hydraulic pump that is airdriven by a variable-pitch propeller.At aircraft speeds above 130

knots, it delivers approximately11.3 gallons/minute (43 liters/minute) at 2,140 psi. On theground, the RAT can be deployedand retracted with the RAT groundmanual switch, located in the rightwheelwell. An onboard RAT check-out module provides verification ofthe operating condition.

Ram A ir Turbine

EAT doc 51 Issue 1



54/167


Controls, Indicators, andCautions

The main hydraulic control panel is

located on the left side of theoverhead panel. The RAT controlswitch is in the center of theoverhead panel. The panels in-clude controls and indicator statuslights. On the r ight side panel arethree switches that control eachsystem isolation valve. On thesame panel is the PTU switch,which allows PTU operation on theground if the left system is unpow-ered.

Control switches for theengine-driven and electric-motorpumps have ON indicators that areilluminated when the switches are

on. Each system has low pressureand reservoir low quantity/pres-sure amber lights, and each pumphas low pressure and overheatamber lights. Reservoir quantityand system pressure can bedisplayed on the lower screen bypressing the EICAS STATUSswitch. The engine fire switches onthe control stand operate the EDPsupply shutoff valves and the EDPdepressurization solenoids.

One caution item, low system pressurethat requires immediate crew awarenessand action, is displayed as follows:

EICAS message.

Two master caution lights.

Amber light illumination on hydraulic

panel.

Aural warning.

Hydraulic Control and Indicators

Advisory items such as low systemquantity, low pump pressure, pump heat, or RAT unlocked that require cawareness are displayed as follows:

EICAS message.

Amber light illumination on

Hydraulic panel.

On the ground, the ELEC/HYD switcthe EICAS maintenance panel canprovide system pressure, reservoirpressure and quantity, and temperatthe lower display.

EAT doc 51 Issue 1



55/167

LLLLLanding Ganding Ganding Ganding Ganding G

Features

GEAR ACTUATION SYSTEM

Landing gears are retracted by theleft hydraulic system, which con-sists of the left engine drivenpump and one electric pump. If theengine-driven pump is inoperative,the right hydraulic system oper-ates a power transfer unit to re-tract the gear.

NOSEWHEEL STEERING SYSTEM

The 757 nose wheel steeringsystems provide 7 degrees ofrudder pedal steering and 65degrees of steering via the tiller.Hydraulic control consists of thesteering metering valve and steer-

ing actuators. A single-loop cablesystem provides inputs to thesteering metering valve. Abroken-cable compensator isinstalled to prevent a sustainedsteering input if the cable fails.

The 757 incorporates the sameconcept used on the 727, 737, and74 7 to prevent rudder pedalsteering when the airplane is flying(landing gear struts not com-pressed).

757 AND 767 COMPARISON

Nose gear for the 757 and 767 issimilar to that on the 737 but islarger. Both the 757 and 767 havetwo unbraked wheels for the nosegear and four braked wheels foreach main gear.

The 757 and 767 brakes, mainwheels, nose wheels, and tires aredifferent, but certified to the sameregulatory requirements. Mainte-nance procedures for the 757 and767 landing gear systems arenearly identical.

Nose wheel steering systems differsomewhat, with the 767 using dualcable loops to provide input to the

steering metering valve, a differenttechnique to prevent pedal steer-ing with the gear retracted, and adifferent hydraulic source.

Main Landing Gear

Nose Landing Gear

Landing Gear Controls and

Indicators

Landing Gear Alternate

Extension

Nose wheel Steering

Proximity Switch System

Brake System

Wheel and Brake System

Components

Brake Temperature Monitori

System (Option)

Ant iskid System

Autobrake System

EAT doc 51 Issue 1



56/167


Main Landing Gear

The main landing gear incorpo-rates a standard air-oil strut for

shock absorption and to supportthe airplanes weight. Each gear ishydraulically extended and re-tracted and incorporates a hydrau-lically operated main door. Themain gear truck is hydraulicallytilted 9.6 degrees forward axle upto provide air/ground sensing. Themain gear is held up and locked byan uplock hook engaging a rolleron the shock strut. The main gearis held down and locked byovercenter locking of a downlock

link. The main gear door actuatorlocks the main gear door closed.Alternate extension is accom-plished by an electric/ hydraulicsystem that unlocks the main gearand doors to allow free-fall exten-sion. Gear position indication isprovided by a dual proximity switchsystem controlled by the proximityswitch electronic unit (PSEU).

Each gear has four wheels andbrakes on a dual-axle truck. Thebearing-mounted brakes are hy-draulically actuated with antiskid

protection provided.

The main landing gear structureconsists of a shock strut, torsionlinks truck assembly, trunnion link,drag strut, side strut, anddownlock assembly.

The shock strut inner and outercylinders provide standard air-oilshock absorption. The strut isserviced with dry air or nitrogenthrough a gas-charging valve on

the top of the strut and with oilthrough an oil-charging valve onthe aft side of the strut. Torsionlinks connect the shock strut innerand outer cylinders. The truckassembly consists of a truckbeam, axles, brake rods, and aprotective shield. The truck beamattaches to the bottom of the innercylinder, providing the pivot pointand attach point for the truckassembly.

There are two fittings and jacpads forward and aft on the trbeam. The bearing mountedbrakes are connected to the in

shock strut by brake rods. A ptective shield on the undersidethe truck protects electrical w

The shock strut mounts to aspherical bearing on the landigear support beam. The trunnlink provides the forward mounand hinge point for the strut towing rear spar. The forward spcal bearing pin connection fortrunnion link acts as a structufuse. The drag strut is a

single-piece brace mounted between the trunnion link and thshock strut to provide fore andstructural suppor t for the gearside strut and downlock assemare two-piece links that fold ovcenter to lock the gear in theextended position and give latstructural support . For groundsafety, a pin is inserted in the of the downlock link.

Main Landing Gear

EAT doc 51 Issue 1



57/167


The reaction link transmits lateralloads from the side strut to theairplane structure and is mountedbetween the wing rear spar and an

inboard support link.

Nose Landing Gear

The nose landing gear incorpo-rates a standard air-oil shock strutfor shock absorption and to sup-port the airplanes weight. Thegear is hydraulically retracted,free-falls to extend, and incorpo-rates hydraulically actuated for-ward doors for an aerodynamicseal. The gear is locked in both the

extended and retracted position byovercenter locking of the locklinks, which are hydraulicallyactuated and aided by a pair ofbungee springs. Alternate exten-sion is accomplished by an elec-tric/ hydraulic system that unlocksthe nose gear doors and allowsfree-fall extension. Gear positionindication is provided by a dualproximity switch system controlledby the PSEU.

Hydraulically powered nose wheelsteering for ground directionalcontrol is provided with tiller or

rudder control. Friction pads breakthe nose wheels on retraction.

The nose gear structure consistsof a shock strut, torsion links, dragbrace, and lock links.

The shock strut inner and outercylinders provide standard air-oilshock absorption. The strut isserviced with dry air or nitrogenthrough a gas-charging valve ontop of the strut and with oil

through an oil-charging valve onthe aft side of the strut. Centeringcams inside the shock strut centerthe gear when extended. Torsionlinks connect the inner and outercylinders, preventing their freerotation and providing a path fornose wheel steering. Forward andaft tow fittings attach to lugs onthe lower inner cylinder. Asingle-piece axle is keyed into aforging on the lower inner cylinder,

providing mounting for the twonose gear wheels.

The shock strut is trunnion

mounted in the nose gear whewell and is supported by atwo-piece folding drag brace. Tupper drag brace is trunnionmounted to wheelwell structurThe lower brace attaches to aforging on the shock strut outecylinder. The drag brace is helocked in both the extended aretracted positions by overcenlocking of lock links; the forwalink is attached to the apex of drag brace, and the aft link to

fitting on the aft nose wheelwebulkhead. Bungee springs andhydraulic actuator provideovercenter locking of the locklinks, which are responsible folocking the gear in the extendand retracted positions.

Nose Landing Gear

EAT doc 51 Issue 1



58/167


Landing Gear Controls andIndicators

A three-posi tion (UP, OFF, DN)

landing gear lever, located on theP3-1 panel, is used to controllanding gear extension and retrac-tion. A lock solenoid in the landinggear lever prevents moving thelever to the UP position when theairplane is on the ground. A lockoverride button is provided. Aguarded ALTN GEAR EXTENSIONswitch controls an electricmotor-hydraulic system that un-locks the main and nose geardoors and gear to allow free-fall

extension.

Position indicators above thelanding gear lever include threegreen gear down and locked lights,an amber gear door open light,

and an amber gear disagreementlight.

Either the captains or the firstofficers brake pedals operateeight hydraulic brake assemblies.A rotary selector switch on the P1-3 panel controls the auto brakesystem. An amber light above theswitch indicates a disarm conditionin the auto brake system. A gaugeon the P3-1 panel indicates brakepressure. Parking brakes are set

by depressing the brake pedalsand pulling a handle on the P1 0quadrant stand.

An amber light forward of thehandle provides indication ofparking brake operation. A resbrakes switch on the P1 -3 pa

isolates the right hydraulic sysac motor pump to the brakes. amber BRAKE SOURCE light the P1-3 panel indicates loss normal and alternate hydraulicbrake source. Optional thermocouple devices on the brakesprovide brake temperature senfor display on the status pageEICAS.

An amber light on the P5 paneindicates antiskid faults. All am

lights have associated EICASmessages.

The rudder pedals permit noswheel steering up to 7 degreeor right, and this may be extento 65 degrees left or r ight by uof the steering tiller on thecaptains auxiliary panel.

Landing Gear Controls and Indicators

EAT doc 51 Issue 1



59/167


Landing Gear AlternateExtension

An alternate extension system is

provided as a backup to the nor-mal landing gear extension sys-tem. The alter nate extensionsystem also opens the landinggear door for ground maintenance.

The alternate gear extensionswitch located below the landinggear lever is actuated to energizea dedicated power pack. Theelectric motor operated hydraulicpump provides pressure to actua-tors for all three gears. These

actuators sequentially operatedoor safety valves to direct thedoor actuator hydraulic fluid toreturn and mechanically unlockdoor actuators and gear up locks.The doors then freely open, andthe gears open by gravity to thedown and locked position. Thealternate extend power pack isthen shut down automatically by apressure switch. All landing geardoors remain open after an alter-

nate gear extension because thedoor safety valves are in theunsafe position.

After alternate gear extension, thelanding gear doors close when thelanding gear lever is moved to theUP position and the gear correctlyretracts with the normal system.

Ground opening of the landinggear doors is commanded by twoALL DOORS OPEN switcheslocated on the P72 panel, acces-sible on the ground aft of the r ightwheelwell. Operation of bothswitches commands the alternate

extend power pack to energize.The actuator operation describedpreviously occurs, placing allsafety valves to the safe positionand opening all gear doors. A redwarning light in each wheelwellilluminates to annunciate an un-safe condition of a safety valve.The red warning lights for the maingear wheel wells are tested beforeentering a wheelwell by operatingthe MLG DR UNSAFE LIGHT

switch on the P72 panel. Operthe NOSE GEAR DOOR UNSALIGHT PRESS TO TEST switclocated on the P62 panel on th

nose gear strut tests the redwarning light for the nose geawheelwell.

Closing the landing gear doorsthe ground requires pressure the left hydraulic system. The landing gear doors are com-manded closed by operating tDOOR CLOSE switch on the Ppanel. Operating the DOORCLOSE switch on the P62 panon the nose gear strut closes

nose landing gear doors. Thesswitches electrically commandhydraulic pressure to door releinterlock actuators, which relethe latching mechanisms, allowsprings to reset the system linages and door safety valves. Lhydraulic pressure then closesdoors.

Landing Gear Alternate Extension

EAT doc 51 Issue 1



60/167


Nose Wheel Steering

Nose wheel steering is controlledby a steering tiller located on the

left side of the flight deck or by therudder pedals. The tiller providesfor turns up to 65 degrees left orright of center. The rudder pedalsgive 7 degrees left or right.

Whether the steering command isfrom the tiller or rudder system(pedals or autopilot rudder rollout),the command signal is transmittedby cables to a hydraulic meteringvalve located on the nose gear.

The metering valve directs hylic pressure from the left systetwo steering actuators to steenose gear wheels.

Internal centering cams in thenose gear shock strut center twheels when the strut is extenafter takeoff, and keep the gecentered when it is retracted aunpressurized during flight.

The steering components incltwo sets of control cables (tilleand piston position), two steeactuators, steering collar, steemetering valve, summing mec

nism and broken cable competor, rudder pedal steering intenect mechanism, torque limiteand a steering tiller.

Nose Wheel Steering

EAT doc 51 Issue 1



61/167


Proximity Switch System

The proximity switch systemprovides position sensing for

landing gear, doors, and thrustreversers. The system consists ofsensors mounted throughout theairplane that sense the proximityof targets and provide positionsignals to the PSEU.

The PSEU is a digital control unitlocated in the main equipmentcenter. It receives signals fromproximity sensors and microswitches; the signals are pro-cessed by software logic that

operates relays, lights, and EICASannunciators. The PSEU alsoincorporates built-in test equip-ment (BITE) to provide in-flightfault detection with storage innonvolatile memory and on-groundtesting of the system.

Air/ground relays transfer variousairplane system control circuitsfrom ground to air mode and fromair to ground mode. The relays are

controlled by the PSEU usinginputs from the main gear truck tiltproximity sensors, the nose gearcompressed proximity sensors,and truck positioner shuttle valvepressure switches.

Two sensors on each main geartruck provide dual system truck tiltinputs to the PSEU. Two sensorson the nose gear strut provide dualnose gear strut compressioninputs to the PSEU.

The sensor inputs are processedin the PSEU logic to drive air/ground rel

51695947 B 757 General Familiarisation in Maintenance

Documents

Transcript of 51695947 B 757 General Familiarisation in Maintenance