B-757-General-Familiarization for Maintenance and Pilots

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

SECTION TITLE ATA

1 Introduction 6, 9, 122 Structures 51, 55, 573 Equipment Centers 11, 20, 23, 394 Flight Deck 11, 25, 33, 35, 395 EICAS 316 Electrical Power 247 Fuel System 288 Auxiliary Power Unit 499 Power Plant Rolls-Royce 71-80

10 Power Plant Pratt&Whitney 71-8011 Hydraulics 2912 Landing Gear 3213 Flight Controls 2714 Environmental Systems 21, 3615 Ice and Rain Protection 3016 Fire Protection 2617 Cabin Systems and Lighting 25, 33, 35,

38, 52, 5618 Cargo Systems 25, 33, 5219 Communications 2320 Indicating and Recording 3121 Navigation 3422 Autoflight 22

Glossary

Table of Contents

StructuresStructuresStructuresStructuresStructures

Features

BASIC STRUCTURALDESCRIPTION

The 757 is metal low-wing mono-plane with full cantilever wing andtail 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 International Air TransportAssociation 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 tothat 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 advancedcomposites extensively in similarapplications such as control sur-faces (carbon) and secondaryfairings (carbon, aramid, fiberglasshybrids).

• Fuselage Reference Diagram

• Fuselage Materials

• Wing

• Wing Center Section

• Horizontal Stabilizer

• Vertical Stabilizer

EAT doc 51 Issue 1Revision 00 date 01/06


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 equipmentbay 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; andhorizontal 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 andcircumferential 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 supported by the frames. Thefuselage is designed to withstandinternal pressure and externallyapplied loads from flight and

ground operating conditions.

The radome forward of the forwardpressure bulkhead is hinged at thetop and made of fiberglass skinsand honeycomb.

The flight deck has three wind-shields on each side, numberedsequentially from forward to aft.The No. 1 windshields are flat andset into forged titanium frames.The No. 2 and 3 windshields arecurved and set into forged alumi-num frames. The No. 2 windshieldscan be opened. Passenger win-dows are made frommoisture-resistant acrylic materialand are mounted in one-piecealuminum forgings.

Fuselage Reference Diagram



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 andlavatories 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 par ts areanodized or alodined and primedwith corrosion-inhibiting primer indetail. In addition, detail partslocated beneath the passengerfloor receive a coat of whiteenamel. Fuselage parts in thelower lobe that are in contact withthe skin or are on the exterior aresealed on contact surfaces.Water-displacing corrosion inhibi-tors are applied to the interiorfuselage structure and to selectedareas of the exterior after allfinishing and sealing. The lowerlobe uses a drainage systemconsisting of drain holes and pathsthrough the structure to permitliquids to reach numerous exter-nally serviceable pressurized drainvalves mounted on the bottomcenterline of the fuselage.

757 Structures



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, rudder,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-strength 350°F (177°C) curecarbon-epoxy pre-impregnated(prepreg) raw material is used forthe majority of the components,and 250°F (121 °C) cure prepregfor the majority of the secondaryhybrid components. Large surfacepanels use honeycomb sandwichconstruction with solid laminateedge bands for attachment tosuppor ting structure. Nonmetallicaramid honeycomb core is sur-faced with composite face skinstailored to provide minimum weightmaximum stiffness components.To prevent galvanic corrosion toaluminum components in contactwith carbon materials, specialprotective systems are used.Fiberglass or aramid plies areco-cured to the carbon contactsurface.

Each aluminum component isanodized, primed, and enameledindividually. An isolating sealant ison all contact surfaces at assem-bly and on all fasteners. The fit-tings and attachments are filletsealed around their peripheries.Corrosion-resistant steel or tita-nium fasteners are used exclu-sively with carbon components.

TITANIUM

Titanium use has greatly in-creased. Titanium alloy forgingsare used in the main landing gearsupport structure and for variousfuselage and nacelle strut fittings.In addition, titanium is used forhigh-pressure tubing and ductingand for firewalls, door thresholds,and scuff plates.Large quantities of coated titaniumfasteners (treated to preventcorrosion) are used, includingsome new types especially devel-oped for use in composite struc-ture.

Composite Material Usage



Wing

The wing surfaces develop aerody-namic forces that support theairplane 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 measuredparallel 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 ribs. 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 platesjoin the skin panels. Upper andlower aluminum stringersstrengthen the skin panels. Thelanding gear is supported by thelanding 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 canti-levered forward from the front sparand is made of aluminum ribs andskin panels. The leading-edgeslats attach to the leading edge.The trailing edge is cantileveredaft from the rear spar and sup-ports the flaps, aileron, and spoil-ers. The wingtip is an aerodynamicfairing covering the outboard endsof the wing. Navigation lightsattach to each wingtip.

The wing, outboard of theside-of-body rib, has access holesin the lower surface between ribs.Similarly, the wing center sectionhas a single hole just to the rightof the keel beam and one accessopening in each of the threespanwise beams. The dry bay overeach engine has four accesspanels to the wing tank, and thetwo ribs immediately outboard ofthe side-of-body splice have ac-cess openings.

Wing Structure



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 anconsists 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 the wingcenter section with the front andrear splice fittings, lower side-of-body splice.. The spar splicefittings are vertically mounted tee-sections machined from the alu-minium alloy. The lower side side-of-body splice is double-shear skinsplice. The splice is an aluminiumchord . The left and right mainsections are joined by more than400 bolts per joint to the wingcenter section to form a unit.

Wing Centre Section



Horizontal Stabilizer

The horizontal stabilizer hassimilar left and right outboardsections 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 ribsjoin the auxiliary 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 therear 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 panels,spars, and ribs. Three actuatorsmove each elevator on eighthinges.

Access to the center sectiontorque box is through the frontspar closeout panel between eachrib. Each closeout panel has twoinspection holes. Inspection of theoutboard main torque box isthrough holes in the rear spar. Thelower surface of the trailing edgehas hinged doors and removablepanels for access to elevatorhinges, actuators, control linkage,hydraulic lines, and wire bundles.Behind the removable leading edgeare inspection holes in the auxil-iary spar. The stabilizer tip isremovable to expose the tip rib.The tip rib has inspection holes toview the outboard ends of thehorizontal stabilizer.

Horizontal Stabilizer Structure



Vertical Stabilizer

The main structural components ofthe ver tical stabilizer are theforward 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 maintorque 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 ver ticalstabilizer 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 isaramid-fiberglass material.The ver tical 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 givesaccess to inspection holes in theauxiliary spar. Removing sectionsof the fin tip allows access to theVOR antenna. Removable panelsin the upper rear spar allow view-ing inside the aft torque box.Removable panels access therudder hinges. Rudder controlsand actuators are accessiblethrough doors in the trailing edgeleft side. The forward torque box isinspected through access panelsand openings from the aft torquebox.

Vertical Stabilizer Structure


Flight DeckFlight DeckFlight DeckFlight DeckFlight Deck

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 flightdecks 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 signalsaler t the flight crew to conditionsrequiring their attention. Thesystem was simplified by minimiz-ing the number of different auralaler ts, 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 acarefully 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 Switches

• Lighting

• Crew Seats

• Windows

• Crew Oxygen system



Design

The design of the 757/767 flightdeck is the result of a long andcarefully planned program ofresearch and development. Theobjective was to provide a flightdeck that would meet the needs ofairline flight crews through the1990’s 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 iner tial referencesystem.Advanced systems monitoring.Built-in test equipment (BITE).

The Boeing flight deck designphilosophy provides enhancedsafety and productivity throughimproved crew comfort and perfor-mance and optimized workload.

Crew comfort is improved byproviding more comfortable anddurable seats, lower noise levels,more efficient air-conditioning, andbetter internal and external vision.

Crew performance is improved asa result of modern-technologyelectronic flight instruments fororientation, a flight managementsystem for airplane navigation andperformance optimization, a cen-tralized crew alerting system, anduncluttered instrument panels.

An optimum crew workload level isneither so high as to cause over-work, nor so low as to causelapses in attention. System designuses simplification, automation,and redundancy to provide asimple man-machine interfacewhile maintaining the proper crewemphasis on flightpath control.

Flight Deck Panels



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. Crewcomfor t and reduced workload areattained by providing adjustableseats, 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 alerts, 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 deckincludes a double crew seat, aflushing toilet, and a rigid cargobarrier.

MainInstrument Panel Arrangement



Lighting Control Panel

Captain’s and First Officers Panel

Glearshield Panel

Main Panel Center

Main Instrument Panel



Overhaed Panel

1. Inertial reference system (IRS) control and display2. Yaw damper3. Hydraulics4. 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 oxygen11. Ram air turbine, engine start12. Fuel system13. Fuel quantity14. Anti-ice15. Windshield wiper

16. Window heath17. Selective calling (SELCAL)18. Passenger sign19. Cabin altitude controls20. Cabin pressure gauges21. Equipment cooling22. Compartment temperature23. Zone temperature/pack control24. Bleed air system

KEY:



Control Stand



Right Side Panel

Basic

Optional



Lighted Pushbutton Switches

Alternate action push-buttonswitches and lights provide controlinputs 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 switchtransfers the switch contacts. Thisswitch does not have switch posi-tion display; however, the lighteddisplay can indicate the position ofa relay or contactor controlled bythe switch.

STATUS/CAUTION DISPLAYS

The status/caution display portionof the switch is a light that displaysa system condition. A legend canbe either a color or black on eithera black or white background.Indication lights use the masterdim and test system power,26.5V-do bright and 12V-do dim.

Lighted Pushbutton Switch (Mechanical)

Alternate action type Momentary type



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 overrideswitch 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



Crew Seats

The captain’s and first officer’sseats move on curved tracks tofacilitate ingress and egress.Armrests, seat backs, and thighpads are manually adjustable forvariations in personnel size. Seatshave lap belts, crotch straps, andshoulder harnesses attached toiner tia reels. The first and secondobservers’ seats are bulkheadmounted and are not adjustable,and fold up when not in use.

Stowage space for suitcases isunder the observer’s seat. A coatcloset is located to the right andaft of the first officer’s seat. Flightkits stow outboard of the pilots’seats.

Miscellaneous equipment andfurnishings in the flight deckinclude crew equipment consoles,glareshield and sunvisors, ash-trays, smoke goggle stowagepockets, hand microphones, head-sets, oxygen masks, observers’panels, removable wastebasket,cupholders, and a spare light bulbholder.

Flight Deck Arranngement



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 significantlyreduce 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



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 regulatorand 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 in acontainer and can be tested with-out removal of the mask.

Overpressure in the oxygen bottlecauses the thermal relief disk torupture, discharging the contentsof the bottle overboard. A thermalrelief indicator on the right side ofthe fuselage shows that this hasoccurred.

Crew Oxygen System


Electrical PElectrical PElectrical PElectrical PElectrical Powerowerowerowerower

Features

AC POWER

Primary electrical power for the757 is generated by the integrateddrive generator (IDG) mounted oneach engine or from the generatordriven 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 thebackup 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 cer tified as essen-tially the same. All major equip-ment is identical or similar, as inthe case of the IDG’s, which haveonly minor engine-dependentdifferences.

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 thesame for the two airplanes.



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. Eachgenerator 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 ofthe essential ac loads in the air-plane. Each bus is divided intoindependent sections. An ac tiebus provides interconnectionbetween the main buses undercertain conditions. The utilitybuses supply nonessential loadssuch as passenger entertainmentand reading lights. Galley power isalso considered a nonessentialload. Nonessential loads can shedautomatically to protect powersources. The optional APU TRUenergizes the APU starter motor ifthe right main ac bus is energizedduring APU star ting. The APUbattery starts the APU if an APUTRU is not installed.

Electrical System Overview



The ground service bus suppliesboth in-flight and ground loads,including interior lights, batterychargers, and cooling fans. Theground 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 selectedcaptain’s and first officer’s 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 primarysource 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 suppliespower 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 chargersystem provides a dedicatedsource of power for operation ofthe standby and autoland systems.The separate APU battery andbattery charger system providespower for APU starting.

Hydraulic Motor Generator(Option)

For extended-range twin opera-tions (ETOPS), the HMG systemprovides a non-time-limited alter-nate source of ac and do powerafter loss of all generator power inflight. An ac generator suppliescaptain’s flight instrument and leftand right transfer buses. A recti-fied generator output powers thehot battery bus.

Electrical Power



Electrical System Panels

Normal operation of the electricalpower system is performed at theelectrical 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 isnot 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 lightilluminates when the main ac busis de-energized.The latching utility bus switchesprovide manual control of thepower relays connecting utility andgalley buses to the left and rightmain ac buses. The ON legendindicates switch position and ishidden when the switch is in theout position. The amber OFF lightilluminates if the associated utilitybus relay is open.

The momentary generator drivedisconnect switches cause amechanical disconnection betweenthe IDG and the engine. Theswitches are spring loaded to theout position and illuminate with lowoil pressure or high oil tempera-ture.

Electrical System Panels



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


AuxiliarAuxiliarAuxiliarAuxiliarAuxiliar y Py Py Py Py Power Systemower Systemower Systemower Systemower System

Features

OPERABLE ON THE GROUND ORIN FLIGHT

The auxiliary power unit (APU)provides an alternate powersource to support aircraft systemson the ground or in flight. It alsoprovides pneumatic power forenvironmental control system andmain engine start. 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 star ting 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 thesame 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 designand 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.

• Auxiliary Power system

• Indication

• Lubrication System

• Fuel System

• Ignition Starting System

• Pneumatic System

• Electronic Control Unit Input and Output



Auxiliary Power System

The auxiliary power system sup-plies electrical and pneumaticpower 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 theAPU 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 andon before the APU can be started.

The APU is warranted to start upto an altitude of 35,000 feet. It iscapable of supplying 115V-ac,3-phase electrical power up to theservice ceiling of the airplane.Pneumatics is available up to analtitude of 17,500 feet. If bothelectrical and pneumatic demandsare present, the ECU reduces thepneumatic output as necessary toprevent exceeding APU EGT limits.The ECU senses five differentpneumatic modes of operationfrom the airplane pneumatic sys-tems.

The ECU positions the inlet guidevanes (IGV) in response to thesemodes to ensure efficient opera-tion and load compressor surgecontrol.

Auxiliary Power System



APU Systems and Components

APU Installation



Indication

The ECU sends analog signals tothe EICAS computers for percentrpm and EGT and discrete signalsfor fault shutdowns and for somefaulty LRU’s stored in nonvolatilememory (NVM). In addition, the oilquantity transmitter sends ananalog signal of oil level directly tothe EICAS.

APU speed in percent rpm andEGT in °C are displayed on theEICAS STATUS and PERF/APUpages.

The advisory message APU FAULTappears and the FAULT lightilluminates to annunciate a protec-tive shutdown of the APU. Thefault light is also used to showtransit of the APU fuel shutoffvalve.

The white RUN light illuminates onthe APU control panel wheneverthe APU is operating above 95%speed.

An hour meter and optional cyclemeter are located on the P49panel in the aft equipment centerto record operating hours and startcycles, respectively. The APUbattery bus powers the hour meter.When APU speed is greater than95%, the ECU provides a groundto operate the hour meter.

APU Indication



Lubrication System

The APU lubrication system con-sists of an oil supply; a pressuresystem for oiling the bearings,generator, and starter 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 arepositive-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 18,000feet), the low ambient air pressurecould cause oil foaming. Thegearbox pressurization systemprevents this by pressurizing thegearbox with second stage com-pressor air; Pcd2 Componentsinclude a gearbox shutoff valve, ashuttle valve, and a gearbox pres-sure-regulating valve. Operation isautomatic and controlled pneu-matically.

Protective shutdowns that areassociated with the lubricationsystem are low oil pressure (LOP),high oil temperature (HOT), andblocked generator oil filter (GENFILTER).

The faulty units stored in the ECUmemory with respect to the lubri-cation system are LOP SWITCH,DEOIL SOL, HOT SENSOR, andFILTER SWITCH.

APU Lubrication System



Fuel System

The APU fuel system receives fuelfrom the airplane wing tanksthrough 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 controlunit, 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 adjustfuel 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 encircle theAPU combustion chamber. Eachhas six fuel nozzles permanentlyattached. The nozzles and mani-fold are an LRU as an assemblyonly.

APU protective shutdowns that areassociated with the APU fuelsystem include NO FLAME, NOACCEL, SLOW START,OVERTEMP, and OVERSPD. NOACCEL and SLOW START areoften caused by too little fuel,while OVERTEMP and OVERSPDare often caused by excessive fuelflows.

Faulty LRU’s that can be displayedon the face of the ECU with re-spect to the fuel system are FUELCONTROL, FUEL SOL, and FLOWDIV SOL.

APU Fuel System



Ignition/Starting System

The ignition/starting system pro-vides initial APU acceleration andcombustion spark. The systemconsists of the ignition unit, theigniter, and the starter 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 orthe optional 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.

Star t initiation opens the APU airinlet door. Once the door is open,the ECU energizes the APU crankcontactor or optional TRU startrelay, as appropriate, to supplypower to the starter 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



Pneumatic System

The APU is designed to providepneumatic power to the airplanefor 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 IGV’s to the load compressorand is discharged into the airplanepneumatic ducts. The IGV’s areessentially 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 IGV’s 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 afunction of air mass flow in theoutput ducting, and a delta-P flowsensor system is used to signalthe ECU for this purpose. 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 Pcd2 power.

The ECU controls the IGV’s andthe surge valve as a function ofairplane pneumatic demand mode,signals from the ECS, sensorsignals, and the settings onswitches located behind a plate onthe face of the ECU. Theseswitches allow adjustments to ECUsoftware without internal repro-gramming to accommodate vari-able external operational circum-stances. The switches are notdesigned for calibration or onlineadjustments.

Air from the plenum is drawn bythe gearbox-driven fan through thefan isolation valve to cool the oiland the APU compartment. Thefan isolation valve is pneumaticallyopened during APU star t whenPcd2 reaches 7.5 psig. Air thenblows continuously through thecooler and compartment. A ther-mal bypass valve prevents cold oilfrom flowing through the cooler.

APU Pneumatic System



A REVERSE FLOW protectiveshutdown occurs if the compressorsurges. This also protects the APUfrom upstream pneumatic systemfailures (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 byturning 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 rpmsignals 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 monitoringand protective shutdown functions.The requirement to interrogate theECU for fault information is annun-ciated in the flight deck by theAPU FAULT light, by the APUFAULT EICAS advisory message,and by the APU BITE EICASmessage on the ECS/MSG page.Some fault LRU’s do not inhibitAPU operation or cause a protec-tive shutdown. The APU BITEEICAS message appears for only12 of the 24 LRU faults. In general,those faulty LRU’s that manifestthemselves by other indications,such as through a protectiveshutdown or loss of pneumaticoutput, do not cause the message.The instructions for accomplishingthe BITE check are on a placardon the door of the E6 rack.

Electronic Control Unit Input and Outpu



The ECU front panel allows faultrecall, reset, and display by use oftwo light arrays and threeswitches. The five-position rotaryFAULT 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 notinhibited, 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 procedurefor 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 FAULTYUNIT switch (down) to annunciateall faulty units stored in theFAULTY UNIT lamp array. Thelamps illuminate from top to bot-tom, left to right across the array.Faults are not sequenced in theorder sorted. TST OK illuminates ifno faults are stored.

ERASE MEMORY. Clear all faultsstored in the ECU by pushing upon the toggle switch located underthe ERASE MEMORY cover. TheWAIT lamp illuminates while theprocedure is in progress, followedby TST OK.

SELF-TEST. This test is identicalto the pre-start BITE accomplishedwhen the APU switch is turned toSTART. WAIT illuminates, followedby any faulty units discovered, thesame as in FAULTY UNIT. If nofaults are detected, TST OK illumi-nates.

Current BTCP 331-200 APU Specifications


HydraulicsHydraulicsHydraulicsHydraulicsHydraulics

Features

TRIPLE REDUNDANCY

Three functionally independent,fulltime, 3, 000-psi systems pro-vide hydraulic power for fullypowered 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 identicalengine 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 nearlyall 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 twodistinct 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 enginedriven 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



Hydraulic Power Distribution

The three hydraulic systems-left,right, and center-are powered by atotal 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 lossof 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 bythe PTU.

The center system has twoACMP’s 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 synthetictype IV fluid (BMS 3-11).

Two hydraulic pumps that aredriven from independent powersources normally power eachsystem. Distribution of pressurefrom the three systems is suchthat the failure of one system willnot result in loss of any flightcontrol functions, and the airplanecan be safely operated in the eventof loss of two hydraulic systems.An emergency hydraulic pump(RAT) provides flight control opera-tion in the event of dual enginefailure.

A central fill point facilitates fluidservicing of all three systems.Reservoir pressurization is ob-tained from the airplane pneumaticsystem and is available wheneverthe pneumatic ducts are pressur-ized. External hydraulic power canbe connected to each system.

Hydraulic Power Distribution



The three systems are color-codedfor easy identification of tubingand components. The left systemis coded red, the center systemblue, 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 ACMP’s, 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 tailfor ground maintenance only.



System Components

The EDP’s 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 islocated in the fluid supply line tothe pump and closes when theengine fire switch is pulled.

The ACMP’s 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 approxi-mately 6.7 gallons/minute (25liters/minute) at 2,850 psi and 9.2gallons/minute (35 liters/minute) at2,000 psi. On-off control of eachunit is provided in the flight deck,and the pumps are on continuallyduring normal operations. Whenthe pump is operating, positivecase drain flow is maintained forcooling and lubrication of thehydraulic pump and electric motor.

The PTU is a fixed-displacementpump driven by afixed-displacement hydraulicmotor. The pump delivers up to 22gallons/minute (83 liters/ minute)at 2200 psi pressure. The hydraulicpump automatically powers thelanding gear and flap/slats actua-tion subsystems if the left systemis not operating and the right EDPis operating.

Hydraulic System Component Location



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 pilot’s overheadpanel for manual deployment atthe pilots’ discretion. 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 asthe 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 130knots, 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 Air Turbine



Controls, Indicators, andCautions

The main hydraulic control panel islocated 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 right 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 areon. 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 over-heat, or RAT unlocked that require crewawareness are displayed as follows:

• EICAS message.• Amber light illumination on• Hydraulic panel.

On the ground, the ELEC/HYD switch onthe EICAS maintenance panel canprovide system pressure, reservoirpressure and quantity, and temperature onthe lower display.


LLLLLanding Gearanding Gearanding Gearanding Gearanding Gear

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).


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 thesteering 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 Monitoring System (Option)

• Antiskid System

• Autobrake System



Main Landing Gear

The main landing gear incorpo-rates a standard air-oil strut forshock absorption and to supportthe airplane’s 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 downlocklink. 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 antiskidprotection 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 onthe 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 jackingpads forward and aft on the truckbeam. The bearing mountedbrakes are connected to the innershock strut by brake rods. A pro-tective shield on the underside ofthe truck protects electrical wires.

The shock strut mounts to aspherical bearing on the landinggear support beam. The trunnionlink provides the forward mountingand hinge point for the strut to thewing rear spar. The forward spheri-cal bearing pin connection for thetrunnion link acts as a structuralfuse. The drag strut is asingle-piece brace mounted be-tween the trunnion link and theshock strut to provide fore and aftstructural support for the gear. Theside strut and downlock assemblyare two-piece links that fold overcenter to lock the gear in theextended position and give lateralstructural support. For groundsafety, a pin is inserted in the apexof the downlock link.

Main Landing Gear



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

Nose Landing Gear

The nose landing gear incorpo-rates a standard air-oil shock strutfor shock absorption and to sup-port the airplane’s 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 theextended 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 orrudder 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 oilthrough 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 trunnionmounted in the nose gear wheel-well and is supported by atwo-piece folding drag brace. Theupper drag brace is trunnionmounted to wheelwell structure.The lower brace attaches to aforging on the shock strut outercylinder. The drag brace is heldlocked in both the extended andretracted positions by overcenterlocking of lock links; the forwardlink is attached to the apex of thedrag brace, and the aft link to afitting on the aft nose wheelwellbulkhead. Bungee springs and ahydraulic actuator provideovercenter locking of the locklinks, which are responsible forlocking the gear in the extendedand retracted positions.

Nose Landing Gear



Landing Gear Controls andIndicators

A three-position (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-fallextension.

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 captain’s or the firstofficer’s 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 setby depressing the brake pedalsand pulling a handle on the P1 0quadrant stand.

An amber light forward of thehandle provides indication ofparking brake operation. A reservebrakes switch on the P1 -3 panelisolates the right hydraulic systemac motor pump to the brakes. Anamber BRAKE SOURCE light onthe P1-3 panel indicates loss ofnormal and alternate hydraulicbrake source. Optional thermo-couple devices on the brakesprovide brake temperature sensingfor display on the status page ofEICAS.

An amber light on the P5 panelindicates antiskid faults. All amberlights have associated EICASmessages.

The rudder pedals permit nosewheel steering up to 7 degrees leftor right, and this may be extendedto 65 degrees left or right by useof the steering tiller on thecaptain’s auxiliary panel.

Landing Gear Controls and Indicators



Landing Gear AlternateExtension

An alternate extension system isprovided as a backup to the nor-mal landing gear extension sys-tem. The alternate 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. Theseactuators 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 rightwheelwell. Operation of bothswitches commands the alternateextend 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. Operatingthe NOSE GEAR DOOR UNSAFELIGHT PRESS TO TEST switchlocated on the P62 panel on thenose gear strut tests the redwarning light for the nose gearwheelwell.

Closing the landing gear doors onthe ground requires pressure fromthe left hydraulic system. The mainlanding gear doors are com-manded closed by operating theDOOR CLOSE switch on the P72panel. Operating the DOORCLOSE switch on the P62 panelon the nose gear strut closes thenose landing gear doors. Theseswitches electrically commandhydraulic pressure to door releaseinterlock actuators, which releasethe latching mechanisms, allowingsprings to reset the system link-ages and door safety valves. Lefthydraulic pressure then closes thedoors.

Landing Gear Alternate Extension



Nose Wheel Steering

Nose wheel steering is controlledby a steering tiller located on theleft 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 hydrau-lic pressure from the left system totwo steering actuators to steernose gear wheels.

Internal centering cams in thenose gear shock strut center thewheels when the strut is extendedafter takeoff, and keep the gearcentered when it is retracted andunpressurized during flight.

The steering components includetwo sets of control cables (tillerand piston position), two steeringactuators, steering collar, steeringmetering valve, summing mecha-nism and broken cable compensa-tor, rudder pedal steering intercon-nect mechanism, torque limiter,and a steering tiller.

Nose Wheel Steering



Proximity Switch System

The proximity switch systemprovides position sensing forlanding 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 thatoperates 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 arecontrolled 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 relays that control variousair/ground critical systems. PSEUand air/ground relay outputs areprovided to the EICAS computersfor air/ground system fault detec-tion and annunciation.

Proximity Switch System



Brake System

PEDALS

Two sets of brake pedals providedifferential braking capability. Thepedal sets are mechanically inter-connected by linkage and drive thebrake metering valves by left andright cable loops.

PARKING BRAKES

The parking brake mechanismlatches all brake pedals in thedepressed position.

The parking brakes are set by fullydepressing the brake pedals andlifting the parking brake lever tolock the pedals in the depressedposition.

HYDRAULIC CONTROL

The normal brake system is pow-ered by the right hydraulic system.The alternate brake system ispowered by the left hydraulicsystem and is automatically se-lected upon loss of the right hy-draulic system pressure. An accu-mulator in the right (normal) sys-tem is automatically selected whenboth normal and alternate systempressure are lost. Two pressureoperated valves select the appro-priate brake pressure source. Areserve braking system is includedin the normal brake system.

The normal brake metering valves(BMV) control brake pressurethrough the autobrake shuttlevalves to the normal antiskidvalves.

The alternate BMV’s meter hydrau-lic system pressure directly to thealternate antiskid valves.

Landing gear up line hydraulicpressure is ported to despin actua-tors on the alternate BMV’s to stopwheel rotation during landing gearretraction.

Hydraulic pressure from the BMV’s(normal or alternate) passesthrough the antiskid valves, thenthrough a shuttle valve and finallyto the wheel brake assemblies.

An optional brake temperaturesensor is mounted in each brakehousing. The brake temperaturemonitor collects and processes thetemperature signals for display onEICAS.

Brake System Diagam



Wheel and Brake SystemComponents

Brake pedal linkage and forwardcable quadrants are located be-neath the flight deck floor andaccessed through the access doorforward of the nose wheelwell.

The brake cables run below theflight deck floor, along the ceilingof the forward cargo compartment,and terminate at the BMV cablequadrants in the right and leftwheel wells.

The brake assemblies are locatedwith each main landing gear wheel.The brake temperature sensorsare mounted in each brake hous-ing. The brake temperature-moni-toring unit is located in the mainequipment center.

The brake accumulator is locatedon the keel beam in the rightwheelwell. The accumulator servic-ing area is in the aft wing to bodyfairing behind the left wheelwell.

The parking brake shut-off valve islocated on the forward wall of theright wheelwell.

The accumulator isolation valve(AIV) and the alternate brakeselector valves (ABSV) are locatedon the forward wall of the rightwheelwell.

Wheel and Brake System Components



Brake Temperature MonitoringSystem (Option)

The monitoring system indicatesindividual brake temperature byproviding eight color coded num-bers in boxes on the EICAS statuspage and a white brake tempera-ture light on the P 3-1 panel. Thebrake temperature monitor unitprovides the capability for BITEtest of the system.

A brake temperature sensor isinstalled in each of the eightbrakes to supply the monitor unitwith a voltage input that is propor-tional to break temperature.

The monitor unit is located on theE5- rack in the main equipmentcenter. It has eight red light emit-ting diodes (LED) to indicatesensor faults, one red LED toindicate a faulted monitor unit, anda rotary test switch,

Brake Temperatuure Monitoring System (Option)



Antiskid System

The antiskid system monitorswheel deceleration and providesbrake release to achieve optimumbraking action under varyingbraking conditions.The antiskid system uses wheelspeed transducer input to applyand release hydraulic brake pres-sure. Four-wheel control cardswithin the antiskid/autobrakecontrol unit control the antiskidfunction. Each card controls a fore/aft pair of wheels through indi-vidual wheel circuits. Airplaneground speed from the inertialreference units (IRU) is provided towheel cards for hydroplane touch-down protection.

PRIMARY ANTISKID FUNCTION

Primary antiskid control is pro-vided on an individual wheel basisfor the normal brake system. In thealternate antiskid system, sepa-rate alternate antiskid valvescontrol laterally paired wheels.

These are operated by signalsfrom the same transducers to thesame control circuits as the nor-mal system but using separatedrivers. Instantaneous wheelvelocity is continuously comparedto a velocity reference value, andthis difference represents the errorsignal. Thus, a wheel deceleratingfaster than the reference would bedetected as entering a skid condi-tion.

SECONDARY ANTISKIDFUNCTIONS

The secondary antiskid functionsinclude locked wheel protection,touchdown hydroplane protection,and gear retract braking.

The purpose of the locked wheelprotection is to allow brake releaseon an individual wheel if a pairedwheel detects a significant slow-down relative to its own speed.

The purpose of the touchdownhydroplane protection system is toensure that the rear wheels haveno pressure applied at touchdownand that the rear wheel brakepressure is released if hydroplan-ing occurs during ground roll.Protection for forward wheels isthrough the locked wheel protec-tion function.

The gear retract braking functioninhibits antiskid control to thealternate antiskid valves for 12.5seconds during landing gearretraction to allow alternate brakepressure to lock the wheels.

FAULT ANNUNCIATION

Faults are annunciated by meansof the amber ANTISKID light andthe EICAS displays on advisory,status, and maintenance levels.

Antiskid System (Simplifed)



Autobrake System

The autobrake system automati-cally applies and controls breakpressure to achieve the airplanerate of deceleration selected bythe flight crew.

The antiskid/autobrake control unitoperates an autobrake module toprovide metered pressure to thebrakes through the normal antiskidvalves. Brake pressure variesaccording to the rate of airplanedeceleration selected and theactual rate of deceleration ob-tained through braking, thrustreversers, and ground speedbrakeoperation.

The autobrake selector switchpowers the autobrake micropro-cessor card in the antiskid/autobrake control unit. It providesfor the selection of desired decel-eration rate. When arming require-ments are satisfied, the selectorswitch receives latching powerfrom the autobrake card. Theswitch panel located on P1-3 alsocontains an amber AUTOBRAKESlight that indicates loss ofautobrake function when illumi-nated.Air/ground mode is provided byrelays that are positioned by maingear truck tilt. The signals areused in the arming and activationof the landing autobrake functions.

The antiskid cards provide wheelspeed inputs to the autobrakecard. The autobrake card providesto the autobrake module the nec-essary signals to achieve therequested deceleration rate.

The BITE card and display cardperform self-test and fault identifi-cation. This information is stored inthe nonvolatile memory of theantiskid/ autobrake control unit,results in the AUTOBRAKE amberlight illuminating, and is sent toEICAS computers for display onthe advisory and maintenancepages.

Autobrake System (Simplified)


Flight ControlsFlight ControlsFlight ControlsFlight ControlsFlight Controls

Features

PRIMARY FLIGHT CONTROLS

All primary control surfaces arehydraulically powered. The eleva-tors and rudder are powered bythree hydraulic systems, and eachaileron is powered by two hydraulicsystems.

SECONDARY FLIGHT CONTROLS

Secondary flight controls include thespoilers and speedbrakes, horizontalstabilizer, leading-edge slats, andtrailing edge flaps. Each spoiler ispowered by a single hydraulic systemand is electrically commanded.Leading edge slats and trailing edgeflaps are mechanically controlled andhydraulically powered.

EFFECTIVE, EFFICIENT HIGH-LIFTDEVICES

The high-lift system consists offull-span leading edge slats andfour trailing edge flaps that arehydraulically powered and com-manded by cables with electricalpower backup. Power is transmit-ted from separate power units tothe surfaces by torque tubes.Automatic flap load relief is pro-vided.

STABILIZER TRIM SYSTEM

A hydraulically powered movablestabilizer that is electronicallycommanded provides pitch trim.

Roll and yaw trim are electricallypowered and commanded.

AUTOPILOT, FEEL FORCES, ANDCONTROL SYSTEM ELECTRONICUNIT

Autopilot control is provided bythree parallel hydraulicservoactuators in pitch, yaw, androll. Feel forces are provided byhydraulic actuators and springs inpitch and by springs in roll and

yaw.Control system electronic func-tions are centralized in dual con-trol system electronic units.


Flight deck indications and con-trols for both airplanes are nearlyidentical. Differences in airplanesize, aerodynamics, and missionrequirements produce some differ-ences in flight controls.

Pitch control-The 757 uses con-tinuous elevator actuatorforce-fight (out of rig) monitors,whereas the 767 relies on periodicground test and pilot tactile detec-tion. The 757 actuators havepressure reducers, and eachelevator is commanded by a singleload path linkage with centeringsprings; the 767 uses a linkagebacked up by a slave cable.

Roll control-The 767 employs bothinboard and outboard ailerons withboost actuators for the controlcables. The outboard ailerons arelocked out for high-speed flight,and the inboard ailerons aredrooped to supplement the flaps.The 757 has one aileron on eachwing and no cable boost. The 767uses all 12 spoilers in flight; theinboard spoilers have an emer-gency evacuation system overrideactuator. On the 757, spoilers 4and 9 are used only on the ground.

Yaw controlThe 757 uses continuous rudderactuation force-fight (out of rig)monitors, in contrast to periodicground test and tactile detectionon the 767.

High lift controlsThe 767 has 12 slats, compared to10 for the 757. The 757 havedouble slotted inboard and out-board flaps; the 767 outboard flapsare single-slotted.

• Overview

• Flight Control Actuators, Servos and Electroncs

• Control System Electronic Unit Interface

• Autoflight Interface

• Aileron System

• Spoiler And Speed Brake System

• Flap Slat System

• Stabilizer Trim System

• Elevator System

• Stall Warning System

• Rudder System



Overview

The primary flight controls are theailerons, which control around thelongitudinal axis (roll); the eleva-tors, which control around thelateral axis (pitch); and the rudder,which controls around the ver ticalaxis (yaw). The secondary flightcontrols are the spoilers andspeedbrakes, horizontal stabilizer,leading-edge slats, and trailingedge flaps.

Flight Control Actuators,Servos, and Electronics

Control cables and associatedlinkages convey mechanical inputsto the power control units (PCU).There is no manual reversion onany primary or secondary flightcontrol. Ailerons have two PCU’sper surface (four total), elevatorshave three PCU’s per surface (sixtotal), and the rudder has threePCU’s total. The distribution ofPCU’s allows triple hydraulicsystem redundancy for the primaryflight control surfaces.

The 12 spoiler panels are operatedby electrically controlled andhydraulically powered PCU’s.

The horizontal stabilizer is drivenby a hydraulically poweredballscrew actuator, which is con-trolled electrically.

Leading edge slats are mechani-cally controlled and hydraulicallypowered using torque tubes androtary actuators. Alternate opera-tion is by electric control and anelectric motor through the sametorque tubes and rotary actuators.

Trailing edge flaps are mechani-cally controlled and hydraulicallypowered using torque tubes andtransmissions. Alternate operationis by electric control and an elec-tric motor through the same torquetubes and transmissions.

Flight Control Actuators, Servos and Electronics



Control System Electronic UnitInterface

The control system electronic units(CSEU) integrate the electronicflight control functions and providea common power supply. EachCSEU contains modules thatperform the functions describedbelow.

The spoiler control modules pro-cess inputs from control wheel andspeedbrake lever transducers toprovide control signals to spoileractuator servo valves.

The stabilizer trim and elevatorasymmetry modules (SAM) providemanual trim and automatic trimduring autopilot operation. Theyalso provide Mach/speed trimcomputation, control of elevatorasymmetry, and programmedairspeed input to the rudder ratiochanger modules.

The rudder ratio changer moduleschange the rudder authority basedon airspeed.The yaw damper modules provideyaw damping, turn coordination,ride smoothing, and fin gust loadreduction.

Control System Electronic Unit Interface



Autoffight Interface

Three flight control computers(FCC) command dedicated autopi-lot servos to control airplanemovement in roll, pitch, and yaw.These autopilot servos engagedirectly into the aileron/spoiler,elevator, and ruddermechanical-hydraulic controlsystems. The FCC’s also input tothe stabilizer trimelectrical-hydraulic control systemfor long-term pitch trim required tocomplement short-term elevatorpitch inputs.

Control System Electronic Unit Interface



Aileron System

Airplane roll control is provided byone aileron on each wing.

The ailerons are controlled byeither pilot’s control wheel. Move-ments of the control wheels aretransmitted to forward quadrants,which are interconnected by a busforce limiter rod, providing overridecapability. Each forward quadrantis connected by cables to quad-rants located in the left and rightwheelwells. The two wheelwellquadrants are interconnected by abus rod and an override on thelateral override quandrant in theright wheelwell.

The left wheelwell quadrant in-cludes the feel and centeringmechanism and the trim actuator.The trim actuator is commandedelectrically by a pair of springreturned toggle switches locatedon the left side of the controlstand.

Trim operation back drives thecontrol wheels and indicatesposition by index marks on top ofthe control column.

The three autopilot servos arelocated in the wheelwells, two inthe left and one in the right. Theyinput by cranks and connectingrods to the wheelwell quadrants.When operating, the autopilotcontrols the ailerons and backdrives the control wheels to com-mand the spoilers.

The wheelwell quadrants areconnected by cables runningalongside the wing rear spar andterminated by a return pulley onthe breakout quadrant locatedoutboard on the wing.

The breakout quadrant contains anoverride and outputs by connectingrod to the control valve of twoPCU’s on each aileron. Each PCUis powered by a different hydraulicsystem.

The aileron movement rotatestransmitters that provide aileronposition on the lower left of theengine indication and crew alertingsystem (EICAS) status page.

Aileron System



Spoiler and Speedbrake System

SPOILERS

Six spoiler panels on each wing,four outboard and two inboard,provide for secondary roll controlas well as flight and groundspeedbrakes.

The spoilers are commandedelectrically by a fly-by-wire system.Six electronic spoiler controlmodules (SCM) each control a pairof spoiler panels symmetricallylocated about the airplane center-line.

Input commands to the SCM’s forroll control are from transmittersoperated by the movement of thecontrol wheels. The captain’s andfirst officer’s transmitters eachcontain three rotary variable differ-ential transformers (RVDT).

Input commands for speedbrakecontrol are from transmittersoperated by the movement of thespeedbrake lever located on theleft side of the control stand. Threepairs of speedbrake lever linearvariable differential transformers(LVDT) are located inside thecontrol stand.

In response to input commands,the SCM’s output command sig-nals to operate their dedicated pairof spoiler panels.

Each spoiler panel is moved by ahydraulically powered PCU con-taining an electrohydraulic servovalve and a feedback RVDT. Acontrol wheel input results inoperation of the PCU’s on onewing only. A speedbrake lever inputresults in operation of the PCU’son both wings. The electricalcommand from the SCM is sent tothe electrohydraulic servo valve ofthe PCU’s, directing hydraulicpressure to the actuators. Move-ment of the actuator piston rotatesa PCU RVDT, which sends anelectrical feedback signal to theSCM to verify execution of thecommand.

The spoiler control system in-cludes redundancy to allow thesystem to operate with an elec-tronic failure. Failures are moni-tored by the SCM’s and are indi-cated by amber light annunciationon the P5 panel and EICAS advi-sory and maintenance messages.

Spoiler and Speedbrake System



AUTO SPEEDBRAKE

An electrical actuator located inthe control stand drives thespeedbrake lever for autospeedbrake deployment and stow-age.

If the speedbrake lever is moved tothe ARMED position before land-ing, upon touchdown the autospeedbrake deploys when boththrust levers are at the idle posi-tion. After landing with thespeedbrake lever in the DOWNdetent, actuation of either thrustreverser deploys the autospeedbrake.

Stowage of the auto speedbrakeoccurs when either thrust lever isadvanced or the airplane is in theair.

A failure of the auto speedbrake isindicated by an amber light illumi-nating and an advisory EICASmessage.

Spoiler System



Flap/Slat System

The flap/slat system consists of 10leading edge slats and four trailingedge flaps. The left hydraulicsystem is the normal source ofpower, with electric motors asbackup.

CONTROL AND INDICATION

Flap lever motion is mechanicallytransmitted to the flap power driveunit (PDU), which is mechanicallyslaved to the slat PDU. Alternateflap/slat control is provided byALTN flap and slat switches and arotary position selector switch.

ALTN switches energize hydraulicbypass valves and arm electricdrive motors. A flap load reliefsystem is provided.

An automatic shutdown of allhigh-lift devices occurs if an asym-metry condition or disagreementcondition is detected. The auto-matic shutdown holds the flapsand slats at the position where theshutdown occurred.

Displays for the flaps and slats onthe center instrument panel in-clude a rotary position indicatorwith a needle for each wing. Amberlights indicate LEADING EDGEand TRAILING EDGE faults. EI-CAS messages are given forcaution, advisory, status, andmaintenance levels.

Flap/Slat Control and Indication



DRIVE SYSTEM

Flaps and slats are each driven bysingle hydraulic motor-poweredtorque tubes. Electric motors drivethe torque tubes in the alternateoperation. Each is driven by twoballscrew actuators. Each slat isdriven by two rotary actuators.

Inboard and outboard trailing-edgeflaps are double slotted and con-trolled by the flap handle throughseven positions.

The leading-edge slats have threepositions: takeoff, cruise, andlanding. In takeoff position, theslats remain sealed with the lead-ing edge of the wing. In landingposition the slats separate (gap)from the wing.

If the flaps are commanded to thelanding position (lever at 30) andthe airspeed is too high, a loadrelief system retracts the flaps tothe flaps 25 position to preventoverloading the wing rear spar.

With flaps in the takeoff position,the slats automatically extend fromtakeoff (sealed) to landing position(gapped) when a stall is detected.The slats are automatically resetto the sealed position when theangle of attack is reduced tonormal. Autoslat operation isinhibited when the airspeed isgreater than 200 knots.

Flep Slat Drive System



FLAP/SLAT ELECMTRONIC UNIT

Three interchangeable flap/slatelectronic units (FSEU) on the E5rack perform separate functionscontrolled by mode select inputsdetermined by rack position. TheFSEU receives alternate positionselector switch, air data and stallwarning computer, and positiontransmitter input depending onFSEU function.

The FSEU performs the followingfunctions: normal control monitor-ing, alternate control, alternatecontrol monitoring, autoslat andflap load relief operation, positionindicating, flap/slat positions forother systems, and hydraulicdepressurization control.

The FSEU faceplate containsbuilt-in test equipment (BITE)instructions, push-button to oper-ate interrogation, and light-emittingdiode (LED) display for messages.

System faults are detected by acontinuous monitor BITE systemand entered in memory. A groundtest function is available to identifyLRU components and verify main-tenance action. A connector pro-vides the capability to connect adata bus analyzer (ARINC 429) tothe FSEU data bus to assist main-tenance.

Flap/Slat Electronic Unit



Stabiliser Trim System

The horizontal stabilizer providesfor airplane pitch trim control byvarying the stabilizer angle ofattack.

The stabilizer ballscrew actuatorassembly is powered by two hy-draulic motors with two hydrauli-cally released brakes. Two stabi-lizer trim control modules (STCM)are electrically commanded toprovide hydraulic power to thebrakes and motors.

Trim commands are from themanual electric trim switches oneach pilot’s control wheel, autopi-lot commands from the flightcontrol computers, Mach/speedtrim augmentation commands, orthe alternate electric trim switcheslocated on the control stand.

Stabilizer trim position is displayedon indicators located on the con-trol stand, outboard of the thrustlevers. Two cutout switches, alsolocated on the control stand,control hydraulic shut-off valvessupplying power to the STCM’s forstabilizer trim control.

Moving the control column in thedirection opposite the trim move-ment actuates cutoff switches tostop trim for all modes exceptalternate electric commands.

Stabiliser Trim System



Elevator System

One elevator on each side of theaft fuselage provides for primaryairplane pitch control.

The elevator is controlled by eitherpilot’s control column. The systemallows one pilot to operate theelevator on one side in case of ajam in the control of the other side.A stick shaker is installed on eachcontrol column to warn the pilotsof an approaching airplane stallcondition.

The two column torque tubes areconnected by a column overridethat provides for jam tolerance.Movements of the control columnsare transmitted by cranks andconnecting rods to the two forwardtension regulator quadrants, whichensure that constant control cabletension is maintained.

The control cables are routedseparately to the aft end of thestabilizer compartment: thecaptain’s cables in the cabinceiling and the first officer’s cablesbetween the cabin floor beams.

Control cables are connected totwo aft quadrants each mountedon torque tubes located aft of thestabilizer. The two torque tubes areinterconnected by an override forjam tolerance. An elevator asym-metry limitation device limits theamount of torque tube relativedisplacement as a function ofairspeed in case of jam. Eachtorque tube is connected by cranksand linkages to the control valveon each PCU. An override on theinput of each PCU valve allowsinput to the other PCU’s in case ofa valve jam. The PCU’s of eachelevator are supplied with a hy-draulic system pressure that isnormally reduced to 2,250 poundsper square inch (psi).

A feel and centering unit located inthe aft stabilizer compartmentprovides for mechanical centeringof elevator controls and for vari-able feel force at the control col-umns. A feel actuator receivesvariable hydraulic pressure fromthe feel computer to generate thefeel force. The feel computermeters the variable feel pressureas a function of airspeed andstabilizer position. The stabilizerposition causes the feel and cen-tering unit to shift the neutralposition of the elevator to provideincreased airplane nose-down trimauthority.

Three autopilot servos, eachcommanded by a flight controlcomputer, are connected by cranksand linkages to the aft quadranttorque tubes.

Elevator System



Stall Warning System

The stall warning system has twodigital stall warning computers(SWC) whose function is to calcu-late when the airplane is nearing astall condition and to provide awarning through operation of thestick shakers. Additional functionsof the SWC’s are to provide inputto the windshear detection andguidance system for visual andaural warning annunciation andflight instrument display and to theFSEU for automatic extension ofthe leading-edge slats.

Each SWC operates a separatestick shaker and provides input toa separate FSEU. Both SWC’sinput to the ground proximitywarning computer and the elec-tronic flight instrument system forthe windshear detection and guid-ance system and to the EICAScomputers for fault annunciation.

Inputs to the SWC’s are flap/slatposition and slat movement(FSEU), body pitch angle and rate(iner tial reference system), dualpower supply modules, Mach, trueairspeed, computed airspeed,indicated angle of attack (air datacomputer), and air/ ground sensing(air/ground relays).

Each SWC has a test switch and aBITE display.

Stall Warning System



Rudder System

A single rudder pivoted on thevertical stabilizer provides forprimary airplane yaw control.Either set of rudder pedals inputsto a pair of cables connected to anaft quadrant located at the bottomof the vertical stabilizer rear spar.The aft quadrant is mounted on atorque tube that also receivesinputs from the three directionalautopilot servos and the trimactuator. Rudder trim is electricallycommanded by a trim knob locatedat the aft end of the control stand.

Rudder trim position is shown on atrim indicator located in front ofthe trim knob. A feel and centeringunit is part of the quadrant torquetube. Output of the quadranttorque tube is by two connectingrods to the rudder ratio changermechanism.

The rudder ratio changer mecha-nism varies the output authority ofthe pedals, autopilot servos, ortrim commands as a function ofairspeed. At low airspeed, fullauthority is transmitted to therudder, whereas it is graduallyreduced with increased airspeed.The ratio changer actuator iselectrically commanded by one ofthe two rudder ratio changermodules. Output of the rudder ratiochanger mechanism is by a pri-mary and secondary control pathto the summing mechanism.

The summing mechanism adds theinputs from the rudder ratiochanger mechanism, the yawdamper servos, and the thermalcompensating linkage to output tothe rudder PCU’s. Two yaw damperservos, each commanded by a yawdamper module, provide inputs tothe summing mechanism todampen dutch roll effects and tocoordinate turns. A thermal com-pensating linkage inputs to thesumming mechanism to overcomethe effect produced by the differ-ence in temperature during climband descent. Output of the sum-ming mechanism is by connectingrods to the rudder PCU’s. Thethree rudder PCU’s are identicaland each is powered by a differenthydraulic system.Rudder movement is sensed by atransmitter to provide for rudderposition indication on the lower leftof the status page on EICAS.

Rudder System


Environmental SystemsEnvironmental SystemsEnvironmental SystemsEnvironmental SystemsEnvironmental Systems

Features

ENGINE BLEED AIR SUPPLY

Engine bleed air and pneumaticground carts provide pressure andtemperature controlled air to usersystems.

BUILT-IN TEST EQUIPMENT FORBLEED AIR SYSTEM

A line replaceable unit in the mainequipment center provides faultmonitoring for the bleed air sys-tem.

MINIMUM-DRAG RAM AIR SYSTEM

Cooling air intake is automaticallycontrolled in flight to position raminlet doors and exit louvers forminimum airplane drag.

AUTOMATIC CABIN PRESSURECONTROL WITH MANUAL BACKUP

Regulating the discharge of airfrom the airplane controls pressureinside the cabin. Manual andautomatic controls for this systemare located on the pilot’s overheadpanel.

AIR-CONDITIONING ON GROUNDOR IN FLIGHT

The auxiliary power unit (APU)supplies all the power needed forair-conditioning on the ground andeliminates the need to tie intoground air supplies.

AUTOMATIC CABIN TEMPERATURECONTROL

Each of the three cabin zones inthe airplane has a separate auto-matic temperature control.

AIR DISTRIBUTION ANDRECIRCULATION

The mix manifold combines condi-tioned air from the left and right airconditioning packs with the recir-culated and filtered air from the

two recirculation fans. Flight deckairdistribution is through ducting tovarious floor, shoulder, and wind-shield outlets. Passenger compart-ment air is distributed from the mixmanifold through sidewall risersand overhead ducts to the passen-ger areas, lavatories, and galleys.

LAVATORY AND GALLEYVENTILATION

Exhaust air from the lavatories andgalleys is routed through a networkof ducts, check valves, and fans.The system exhausts into the lowerlobe next to the outflow valve andthen discharges the air overboard.


The 757 and 767 use basically thesame heating, cooling, pressuriza-tion, and air-conditioning packs.

Components for the cabin pressurecontrol system are identical. Be-cause of airplane performancedifferences, the climb schedules forthe two airplanes vary and are pinselectable for each model.

The 757 and 767 electrical/elec-tronic cooling systems are function-ally and operationally similar. Bothsystems use similar fans, ducting,valves, and avionics installations,and both have similar warning andindication systems.

The 757 air distribution and recir-culation systems are similar tothose on the 767 with a few differ-ences. Lavatory and galley ventila-tion systems are essentially identi-cal.

The 757 freighter has an additionalpair of supply fans devoted tocooling flight deck electrical/elec-tronic equipment. This system,which was designed to preventupper deck cargo air from enteringthe flight deck, incorporates thesame type of control and indicationsystem as the 757 main supplyfans.

• Pneumatics

• Environmental Control System

• Air-Conditioning

• Cargo Heating and Lavatory and Galley Venting

• Cabin Pressurization

• Electrical Electronic Equipment Cooling



Pneumatics

PNEUMATIC SYSTEM OPERATION

The pneumatic system suppliescompressed air forair-conditioning, engine starting,hydraulic reservoir pressurization,wing and engine cowl anti-icing,and potable water tank pressuriza-tion. The system automaticallycontrols the temperature andpressure of the air supply withassociated indications on theoverhead panel.

Low- or intermediate-pressure airis automatically supplied as re-quired from a port on the engine.The air flows through a checkvalve, which prevents reverse flowduring high-pressure operation, toa precooler. This air is used duringclimb, cruise, and cer tain holdingconditions.

High-pressure air is automaticallysupplied as required from ports onthe engine.

The high-pressure controllersenses engine case pressure andsupplies servo pressure to controlthe high-pressure shut-off valve.The high-pressure shut-off valvepneumatically modulates to supplyair at 55 ±5 pounds per squareinch (psi) or to shut off thehigh-pressure supply of air.High-pressure air is used duringdescent, low power settings, andcertain holding conditions.

The temperature of the bleed air isregulated to 3800 ±20 °F (1930±11 °C via a precooler that func-tions as a crossflow-type heatexchanger. The precooler usesengine fan air as a heat sink. A fanair modulating valve regulates theamount of cooling air that is sup-plied to the precooler. The valve ispneumatically controlled by a fanair temperature sensor that sensesthe bleed air temperature down-stream of the pressure regulatingand shut-off valve (PRSOV) andsupplies a servo pressure signal tothe fan air modulating valve.

Flow control is regulated via thePRSOV and controller. The PRSOVcontroller is solenoid controlled by28V-DC signal from an alternateaction switchlight on the overheadpanel. The PRSOV is a pneumati-cally actuated valve that regulatesthe bleed air pressure to 45 ±1 psi.The PRSOV controller and systeminclude a temperature- I limitingfunction. A reverse flow controllerprevents reverse flow in the engineair supply system from the airdistribution system.

An overtemperature switch pro-vides a signal to the bleed light onthe P5 overhead panel whenever atemperature of approximately490°F (254 °C) is sensed. Theengine bleed system is com-manded off automatically withoutcrew action.

Engine Bleed Air System Control



Pneumatic Systems



The engine air supply system ismonitored and controlled from thepilots overhead panel. Alternateaction switchlights labeled L ENGand R ENG control the PRSOVand high-pressure controllers.Amber lights, in conjunction withthe caution and warning system,indicate overtemperature condi-tions (BLEED), PRSOV closed(OFF), and high-pressure valve/controller malfunctions (HISTAGE).

PNEUMATIC DISTRIBUTION

An isolation valve isolates air fromthe pneumatic sources. Normal airsupply to each air-conditioningpack is from each respectiveengine. During ground operation,the auxiliary power unit (APU) canbe used to supply the left packwith the valve closed or both packswith it open. The isolation valve ismotor driven and is controlled fromthe pneumatic control panel.

For normal flight operation, theisolation valve is closed. In the

event of one bleed source outcondition, the isolation valve canbe opened to supply both wingthermal anti-ice (TAI) systems andan air-conditioning pack from theremaining bleed air source. Ductpressure indicating is provided inthe flight deck to display the pneu-matic duct pressure on each sideof the isolation valve. The systemprovides air for water and hydraulicreservoir pressurization.

With the APU operation at 95% orabove, the APU can be used as anair supply source. The APU isola-tion valve is controlled by analternate action switchlight. Theswitchlight includes an amberdisagreement light that, along withthe caution and warning system,indicates valve malfunction. Acheck valve prevents flow to theAPU.

An overheat detection system isprovided to detect leakage fromthe bleed air duct system and toprevent overheating the structure.

The overheat detection systemconsists of two independent,continuous sensors installedadjacent to the pneumatic ductsystem. In the event of hot airleaks, impingement on the sensorsactivates the caution light locatedon the pilots overhead panel andwarns the crew of the overheatsituation.

Detection is divided into left andright systems. The left systemdetects the left wing leading-edgecavity, left air-conditioning pack,and APU supply duct area, and theright system detects the right wingleading-edge cavity and right air-conditioning pack.

The air pressure in the ducts oneach side of the pneumatic mani-fold is displayed on a dual indica-tor pressure gauge located on theP5 overhead panel. Duct pressureis also available on the ECS mes-sage page of the engine indicationand crew aler ting system (EICAS).

Pneumatic Distribution System



Environmental Control System

OVERVIEW

An energy-efficient air-conditioningsystem provides passengers andcrew members with conditioned air.Energy efficiency is accomplishedby only requiring that about 50% ofthe total conditioned air volume bederived from engine bleed. Theremainder is provided by filtered,recirculated air. A three-zoneautomatic temperature controlsystem is provided for two maincabin zones and the flight deck.The environmental control system(ECS) incorporates system testand line replaceable unit (LRU)fault-isolation capability.

The lavatory and galley ventilationsystem provides positive ventila-tion of smoke and odors from thelavatories and galleys.

The forward cargo compartment isheated by conduction through thecargo lining and air recirculation.The aft cargo compartment has anelectric heating system.

A cooling system is provided forthe forward and aft electrical/electronic (E/E) equipment coolingsystems to meet FAA and CAArequirements. Filtered air is forcedthrough the forward system by afan. The warm air is removed bythe left air-conditioning recircula-tion fan in the forward E/E equip-ment cooling system and by thelavatory and galley ventilationsystem in the aft E/E equipmentcooling system.

Environmental Control System



ECS CONTROL

The air-conditioning system iscontrolled from the pilots overheadpanel, and each pack has separateautomatic controls. In the automode, each pack temperaturecontroller automatically positionsthe temperature control valve andram air inlet and exit doors. Byproper scheduling, the pack tem-perature controller minimizes ramair drag without exceeding equip-ment maximum temperature limita-tions.

Each pack has independentstandby control (N for normal, Cfor cold, and W for warm). Duringnormal operation, the water extrac-tor temperature is controlled bythe low limit valve to 40 °F (40°C).In warm operation, the tempera-ture is equal to primary heatexchanger outlet temperature,bypassing the air cycle machine.In cold operation the temperatureis maximum cold, and all airpasses through the air cyclemachine.

Alternate action switchlights withINOP/OFF lights control operationof the two recirculation fans andsupply power to the trim air pres-sure regulating and shut-off valve.

Each zone has an automatictemperature control provided bythe zone controller, and no manualcontrol is provided. A trim valvewill close when its respectivetemperature selector is placed tothe OFF position.

In the event that the zone control-ler malfunctions or if the mastertrim air switch is placed in the OFFposition, the system automaticallytransfers temperature control tothe pack controllers. The left packcontrols the flight deck to a fixed75°F (24 °C) and the right packcontrols the passenger cabin to afixed 75 °F (24 °C).

Indicators allow monitoring of packoperations and compartmenttemperatures. Amber lights, inconjunction with EICAS, allowsystem fault monitoring.

ECS Control



The doors are driven by actuatorsthat are controlled by the respec-tive auto pack temperature control-ler. This system cools the bleed airusing crossflow heat exchangerswith ram air as the heat sink. Aseries arrangement of the heatexchangers has the ram air pass-ing through the secondary heatexchanger before the primary.Airflow is induced through the ramair system on the ground or duringlow flow conditions by a fan that ispart of the air cycle machine(ACM).

Cold air is provided by the ACMturbine. The ACM has three rotat-ing impellers consisting of a com-pressor, a turbine, and a fan,which are mounted on a commonshaft. The expansion of the airthrough the turbine cools the airand generates power to drive thecompressor and fan. Compressoroverheat protection is by thermalsensors and a thermal switch.A high-pressure water separationsystem, consisting of a condensor,

water extractor, and reheater,removes moisture from the air. Thecondensor uses crossflow ofturbine outlet air to cool the air,which allows the moisture tocondense into droplets. The waterextractor uses helical swirl vanesto spin the air and allow centrifugalforce to remove the droplets. Thewater is ducted and sprayed intothe ram air upstream of the heatexchangers, increasing the effi-ciency of the ram air coolingsystem. The reheater increasesthe temperature of the air beforeflow into the turbine, increasingthe efficiency of the ACM. Elimi-nating a collector bag makesscheduled maintenance unneces-sary.

The mix manifold mixes the condi-tioned air supply from theair-conditioning packs, whichmakes up about 50% of the totalvolume, with the recirculated,filtered air supplied by the tworecirculation fans, which makes upthe remaining 50%.

Air-Conditioning

CONDITIONED AIR SUPPLY

Bleed air is provided to each air-conditioning pack by the pack flowcontrol valve. The valve is a venturitype, solenoid controlled andpneumaticaliy actuated, and hasthree flow schedules. The flowcontrol valve maintains a predeter-mined airflow schedule as a func-tion of airplane altitude. The highflow schedule automatically in-creases airflow through the valveby 65% above normal airflow in theevent of failure of the other pack orof the recirculation tan on thesame side, or when operating thesystem with the APU or pneumaticground car ts. When operating inthe standby mode the high flowschedule is limited to an increaseof 45% above the normal airflowschedule.

Ram air is regulated through twoheat exchangers by the ram airinlet and exit doors.

Cconditioned Air Supply



The two separate and identicalpack temperature controllerscontrol the temperature of theirrespective air-conditioning packsusing the temperature demandsignal generated by the zonecontroller. The controllers controlthe pack to the high-pressurewater separator temperature, themix manifold temperature, and theACM compressor discharge tem-perature. Using the temperaturecontrol valve and the ram airactuators, the pack temperature isregulated based on the system’stemperature demand.

A pneumaticaliy actuated differen-tial pressure low-limit valve pro-vides protection against icing ofthe condenser. The valve allowshot air to bypass the ACM duringSTBY N operation.

The compressor overheat switchprovides shut-off of the flow con-trol valve whenever an overheat of490°F (254°C) is sensed at thecompressor discharge. A packoutlet temperature of 190°F (88°C)configures the pack to full cold.

The zone control system gener-ates a temperature demand signalbased on the zone that requiresthe most cooling and provides thissignal to each pack controller. Thepack controllers then direct thepack temperature control valve andram air actuators to produce thedemanded temperature. The zonecontroller, by way of the varioustrim air valves, adds heat to thoseremaining zones that require heat.The zone controller limits the zoneduct temperature to between 350and 160°F (21° and 71 °C Eachzone duct is provided with over-heat protection by a 190°F (88°C)thermal switch.

CONDITIONED AIR DISTRIBUTION

Conditioned air from the left packis mixed with trim air and suppliedto the flight deck through ducts. Airis supplied in the flight deckthrough sidewall, windshield, andindividual crew outlets. Air sup-plied to the shoulder outlets canbe heated by electrical heaterscontrolled from the captain’s andfirst officer’s consoles.

Conditioned air from the mixmanifold is mixed with trim air andsupplied to the forward and aftpassenger zones through four riserducts (two per side) to the over-head distribution duct. The over-head distribution duct provides theconditioned air to an outlet in thecenter of the ceiling and tosidewall outlets on both sides ofthe passenger compartment.

Conditioned Air Distribution



Cargo Heating and Lavatory andGalley Venting

Smoke and odors are vented fromthe lavatory and galley areasthrough overhead ducting. Forcedventilation is induced by the ventfans.

The forward and aft cargocompartments have separateclosed-loop heating systems. Airfrom the forward cargo compart-ment is drawn through the inletgrille at the aft end of the compart-ment by the cargo compartmentfan. The air is warmed by heattransfer from the cargo com-partment fan motor heat. Thewarmed air is then distributedthroughout the compartment. Atemperature sensor located nearthe inlet of the cargo heatingdistribution duct automaticallycontrols the compartment fanoperation.

The aft cargo compartment heat-ing system works in a similarmanner except that the heating isprovided by both the fan motorheat and an electric heater in thedistribution duct.

Cargo Heating and Lavatory and Galley Vent Systems



Cabin Pressurization

CABIN PRESSURIZATION SYSTEM

Control is from the pilots overheadpanel with both automatic andmanual modes. Two separate andindependent controllers located inthe main equipment center provideautomatic pressurization control.The controllers have separatepower sources and send output toseparate ac motors on the outflowlow valve.

Both air data computers providethe actual airplane altitude infor-mation to both controllers.

When the engine throttles areadvanced more than 10.5°, mi-croswitches in the throttle quad-rant signal the appropriate control-ler to select the pre-takeoff mode.

Air/ground system 1 is dedicatedfor Auto 1, and system 2 is dedi-cated for Auto 2. Each system provides air/ groundstatus information to the controller,which uses the information todetermine mode of operation.

The cabin pressure is controlled byregulating the discharge of airfrom the airplane via the outflowvalve. The valve is driven by eitherof two separate alternating current(ac) motors during auto mode orby a direct current (dc) motorduring manual mode of operation.

Two positive pressure relief valvesprovide airplane over-pressuriza-tion protection. Two negativepressure relief doors provideprotection against airplane nega-tive pressurization, such as maybe encountered during rapid de-scents. The cabin altitude warningswitch signals a horn to alert thecrew of excessive cabin altitude.

Cabin Pressurization System



CABIN PRESSURIZATION CON-TROL

Automatic pressurization control isprovided in two identical modes bytwo separate controllers. In theautomatic mode, the systemresponds to preflight crew inputfrom the selector panel, from theair/ ground systems and actualairplane altitude provided over theARINC data bus, and from apneumatic signal indicating actualcabin pressure. The only crewaction is the preflight selection ofmode of operation, landing fieldaltitude, and the desired maximumlimit for rate of cabin pressurechange.

Automatic controller switchoveroccurs when the controller sensesexcessive cabin pressure changerate, excessive differential pres-sure (more than 8.8 psi), ac powerloss, or self-test fault.

If both auto modes fail, the manualmode serves as a backup. In themanual mode of operation, theoutflow valve is directly driven andheld in position through input fromthe control panel to the dc motor.

Annunciator lights, in conjunctionwith the EICAS, alert theflightcrew of auto controller failure,excessive cabin altitude (greaterthan 10,000 feet (3000 meters)),and, as an option, low air flowcondition. Indicators monitor thecabin altitude, cabin rate ofchange, and cabin-to-ambientdifferential pressure.

A valve position indicator showsrelative position (open or closed).

During auto control, the cabinpressure auto controller automati-cally operates in different modesand schedules.

In the powerup mode, the systemgoes into a self-test. The groundmode opens the outflow valve. Inthe takeoff mode the outflow valvebegins to close and the controllerautomatically modulates the out-flow valve to prevent a pressurebump. During climb, cruise, anddescent, the system operates inthe proper mode according toprogrammed schedules and takeoffand landing altitudes. In the land-ing mode, the system depressur-izes the airplane. Excessive cabinaltitude (greater than 11,000 feet(3400 meters)) automaticallycloses the outflow valve regardlessof input.

Cabin Pressurisation Control Panel



Electrical/Electronic EquipmentCooling

The E/E equipment cooling systemprovides a supply of cooling air tovarious pieces of E/E equipmentand evacuates the hot exhaust airfrom this equipment.

The equipment cooling system isdivided into a forward and an aftsystem. Each system has eitherblow-through or drawthroughcooling that is controlled automati-cally, with alternate control as abackup. Indications are providedfor potential overheats, actualsmoke within the systems, andinsufficient flow rate.

The forward equipment coolingsystem incorporates an air clean-ing system, a blow-through coolingsystem (supply fans), adraw-through cooling system (leftrecirculation fan), smoke detectionand indicating circuit, low flowdetection and indicating circuits,and overheat detection and indi-

cating circuits. The aft equipmentcooling system incorporates adrawthrough cooling system (lava-tory and galley vent fans). Theforward system also has a smokeclearance and differential coolingcircuit. The systems normallyoperate automatically through theuse of airplane circuitry. Somecomponents are tested automati-cally for proper operation everytime both engines are shut down.The system may be tested manu-ally using the EQUIP COOL testswitch on the P61 panel.

The EICAS computers monitor theequipment cooling system forproper operation and providemessages to indicate malfunctionor normal operation. Indicationsare also provided on the equip-ment cooling control panel or theair-conditioning control panel, andthrough the ground crew callsystem.

Electric/Electronic Equipment Cooling System


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Features

WING THERMAL ANTI-ICING

Directing engine bleed air to threeof the four outboard leading-edgeslats on each wing provides in-flight ice protection.

ENGINE INLET THERMALANTIICING

Engine bleed air is directed to theengine cowl inlet lip to prevent iceformation.

WINDOW HEAT

Flight deck windshields are electri-cally heated to prevent ice and fogbuildup. The flight deck side win-dows are electrically heated foranti-fogging only.

PROBE HEAT

Electric heat is provided for thefour pitot static probes, twoangle-of-attack probes, and onetotal air temperature probe.

WATER AND WASTE HEAT

Electric heating is automaticallysupplied to the water and wastesystems to prevent freezing.

WINDSHIELD WIPERS AND RAINREPELLENT

Rain repellent is used with thewindshield wipers to Improvevisibility during heavy precipita-tion.


The ice and rain protection sys-tems on the 757 and 767 areoperationally the same and differonly in size. Engine bleed air isused to heat the inside of theleading edges of the engine inletcowl and leading edge slats. Thepitot probes on the 757 and thepitot-static probes on the 767 areelectrically heated.

When the 757 passenger -andfreighter airplanes are equippedwith identical engine models, theanti-ice and rain removal systemsare identical.

As an option, the 767 offers pri-mary automatic ice protectioncontrol for the wing and engineinlet ice protection systems. Thesystem uses dual redundantairframe-mounted ice detectors.

The same Boeing rain repellent isused on all Boeing aircraft.

• Engine Cowl Thermal Anti-Icing

• Wing Thermal Anti-Icing

• Electrical Ice and Rain Protection



Engine Cowl Thermal Anti-icing

The engine air inlet cowl isanti-iced by engine bleed air. Acircular integral spray duct insidethe cowl distributes the bleed aircircumferentially through rows ofholes that direct jets of hot airagainst the inside of the inlet cowl.The thermal anti-ice (TAI) airexhausts through a slot in thebottom of the engine inlet (P&W)or into the engine inlet (R-R).

A cowl TAI valve is provided foreach engine. This valve, a pres-sure regulating and shut-off valve(PRSOV), controls bleed air to aregulated pressure in normaloperation.

Control of each engine cowl TAIvalve is through an alternateaction switch located in the flightdeck overhead panel. Integral toeach switch are lights for indicat-ing system status. A high-pressureswitch is provided to indicate valvefailure to regulate for Rolls-Royceengines. For Pratt & Whitneyengines, position switches on thevalve monitor proper valve opera-tion. The failure message is dis-played on the status page of theengine indication and crew alertingsystem (EICAS) display.

Engine Cowl Anti-ice Protection



Wing Thermal Anti-icing

Regulated precooled engine bleedair provides ice protection for threeof the four outboard slats (2, 3,and 4 on the left side and 7, 6, and9 on the right side) on each wing.

The engine bleed air systemsupplies pressure- andtemperature-controlled air to thewing TAI systems. In normal opera-tion, each wing is an independentsystem extracting bleed air fromthe engine on that side. Crossfeedcapability is provided to anti-iceboth wings from a single bleedsource. The wing TAI PRSOVcontrols the system pressure.Downstream of the wing TAIPRSOV, a telescoping duct trans-ports the TAI air from the supplyduct to the TAI spray ducts. Thehot air is then distributed spanwisealong the entire length of the threeslats. The TAI air exhausts throughexit slots located in the lower skinof each slat.

Control of both wing TAI valves isthrough a single alternate actionswitch located in the overheadpanel. An amber valve light foreach wing is provided to indicatethe system’s response to pilotaction. The VALVE lights are ononly if there is disagreement fromthe commanded position. A groundtest function is also incorporated.

A failed valve can be manuallylocked in the closed position.

Wing Anti-Ice Protection



Electrical Ice and Rain Protection

WINDOW HEAT PROTECTION

Two separate and identical threechannel window heat control unitsprovide electrical heating toanti-ice the forward windshieldsand to anti-fog the side wind-shields. The control units crosscontrol the windshields to preventloss of protection for one side.Fail-safe protection is provided toshut down a controller for any ofthe following conditions: inputpower not present, shorted sensor,open sensor, overheat, demandwithout heat, heat without demand,or asymmetrical output. The sys-tem is controlled from the pilot’soverhead panel and can be resetby recycling the respective alter-nate action switch light. A systemtest feature is incorporated. Amberlights in conjunction with thecaution and warning system indi-cate system malfunctions.

The controllers incorporate built-intest to identify faulty LRU(line-replaceable units). A pneu-matic backup anti-fogging systemis provided for the No. 1 wind-shields. The pneumatic systemconsists of two nozzles that blowair continuously over the insidesurface of the No. 1 windshieldswhenever the air-conditioningsystem is operating.

PROBE HEAT PROTECTION

Electric heater protection is pro-vided for the pitot, angle-of -at-tack, and total air temperatureprobes. Heating is controlledautomatically in flight and on theground. The pitot probes switchfrom a high to a low heat conditionwhen the airplane lands. No heatis applied when the airplane is onthe ground with all engines shutdown.

RAIN PROTECTION

Dual-speed electrically operatedwindshield wipers are provided forthe two forward windshields. A rainrepellent solution from a singlesupply is available to indepen-dently apply fluid to each of thetwo forward windshields. Thesystem is controlled from thepilot’s overhead panel.

WATER AND WASTE HEATING

Electrical heating is automaticallysupplied to the water and wastesystems when airplane power ison, decreasing the necessity ofdraining the system for overnightstopovers. The drain masts auto-matically switch from a high to alow heat condition when the air-plane is on the ground. The wastedrains are heated via a heatedgasket. Waterlines are heated withheater tape, and the water tank iscovered with an insulating blanket.

Electrical Ice and Rain Protection


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Features

ENGINE

Each engine has dual-loop over-heat and dual-loop fire detectors.Two fire extinguisher bottles canbe directed to either engine.

AUXILIARY POWER UNIT

The auxiliary power unit (APU) hasa dual loop fire detector abovenear the air inlet and below on theright access door. A single-bottleextinguishing system is provided.

LOWER CARGO COMPART-MENTS

The lower cargo compartmentshave dual smoke detectors thatactivate the fire warning system ifsmoke is detected. There are twoextinguishing bottles locatedforward of the aft cargo compart-ment, and either one or both canbe discharged into either compart-ment.

WING AND BODY DUCT LEAKDETECTION

The pneumatic duct leak detectionsystem is divided into right and leftzones and is designed to notify thecrew of a duct rupture. There is noextinguishing system.


The 757 and 767 engine and APUfire detection and extinguishingsystems are very similar in indica-tion and operation. Wheelwelldetectors and engine fire protec-tion systems differ only in therouting and positioning of sensingelements. Although three differentmanufacturers supply fire protec-tion systems for these aircraft, alloperate in the same manner andtrigger the same indicators on theflight deck.

Cargo compartments for bothairplanes are equipped with smokedetection and fire extinguishingsystems. The smoke detection andfire extinguishing systems on the757 and 767 passenger airplanesare essentially identical in func-tion. In addition, the cargo com-partment warning alarms areidentical on both airplanes.

The use of Halon is universal. It isthe extinguisher of choice for itsability to smother fires withoutdamaging sensitive components.

FREIGHTER

The main deck cargo compartmenton the freighter has a continuousair sampling system for smokedetection, and the lower cargocompartments have the samesystems as the passenger air-plane. There are no fire extinguish-ers for either main deck or lowerdeck cargo. Fire is extinguished bydepressurizing the airplanes,which reduces the oxygen avail-able for combustion.

• Overview

• Engine Fire and Overheat Detection and Warning

• Engine Fire Extinguishing

• APU Fire Detection and Warning

• APU Fire Extinguishing

• Cargo Compartment Fire Detection and Warning

• Cargo Compartment Fire Extinguishing

• Wheelwell Fire and Duct Leak Detection

• Freighter Fire Detection



Overview

The fire protection system consistsof detection systems and extin-guishing systems. A number ofindicators on the flight deck assistin fire and smoke detection andwarning, and several methods areavailable for extinguishing fires.

Each lavatory has a smoke detec-tor alarm in the ceiling and a fireextinguisher above the wastecontainer that is automaticallydischarged when the heat-fusibleplugs melt. There is no indicatorfor this on the flight deck.

Fire Protection Capabilities

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b The freighter has no extinguishers for cargo areas

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Engine Fire and OverheatDetection and Warning

Engine fire and overheat detectionis provided by two independentdual loop detector systems oneach engine. Several temperaturelimits are integrated into the sys-tem, depending on the sensorlocation on the engine.

The logic that controls the systemnormally requires both loops toprovide a warning signal before afire or overheat alarm is triggered.A single-loop signal is indicated onthe engine indication and crewalerting system (EICAS). Thesecond loop signal, if initiated,causes fire or overheat alarmactivation. A fire is indicated by afire bell; illumination of the masterwarning lights, fire discrete warn-ing light, fire handle and fuelcontrol switch lights; and a level“A” warning display on the EICAS.Pressing either master WARNING/CAUTION light switch or pullingthe fire switch silences the fire bell

and resets the master warninglights.

An engine overheat is indicated bya caution aural tone and illumina-tion of the master caution lights.The corresponding engine over-heat light and a caution display onEICAS also illuminate.

The complete engine fire andoverheat detection system can betested before and during flight bysimulating fire and overheat condi-tions. Operation of the twoswitches on the FIRE/ OVHT TESTmodule, located on the pilotscontrol stand (P8), actuates asimulated fire and overheat condi-tion If the element loop and controlunit are operating properly, thealarm devices will be energized.

APU Fire Detection and Warning



Engine Fire Extinguishing

The engine fire extinguisher dis-charge switches are incorporatedin the fire switch handles on thecontrol stand (P8). A two-stepprocess arms and activates theextinguishing system.

Pulling either fire switch handledoes the following:

• Closes the engine fuel supply valves.• Closes the engine bleed valve and isolation valve.• Trips the generator.• Closes the hydraulic supply valve.• Arms the extinguishing system.• Silences the bell and resets the master warning lights.• Closes the thrust reverser isola tion valve.• Closes the APU bleed valve (from left fire switch handle only).

Turning the fire switch handleclockwise or counterclockwisedischarges one fire extinguisherbottle and, if turned the oppositedirection, discharges the secondbottle. The bottles are installed inthe forward portion of the aft cargocompartment. The engine bottledischarges light on the pilotscontrol stand illuminates when thepressure switch on the fire bottleindicates that the extinguishingagent has been discharged.

The extinguisher bottle explosivesquibs are tested using the testswitch on the right side panel(P61). Illumination of the squiblights indicates operational squibs.The bottle pressure switch can betested with a switch actuator onthe bottle.

Engine Fire Extinguishing




A dual-loop fire detection system,consisting of upper and lowerdetector elements mounted in theAPU compartment, provides a firewarning through the central warn-ing system. An APU fire-warningsignal initiates an APU auto shut-down.

Fire detection and warning elec-tronics cards process the detectorsignals that generate warnings.

APU fire detection indication in theflight deck consists of a red an-nunciator light in the APU firehandle, master warning lights,bells, and EICAS display. A redAPU firelight and a horn are alsoexternally mounted in the APUremote control panel (P62) on thenose landing gear.Pressing the master WARNING/CAUTION switch lights or pullingthe fire handle turns off the bell.

The complete APU fire detectionsystem can be tested before andduring flight by simulating fireconditions. Operating the ENG/APU/ CARGO switch on the FIRE/OVHT TEST module located on thecontrol stand (P8) actuates asimulated fire condition. If theelement loop or the control unit isinoperative, the EICAS system willindicate the faulty loop.




APU Fire Extinguishing

The APU fire extinguishing systemconsists of a single fire bottlelocated forward of the firewallbulkhead. The controls for releas-ing the extinguishing agent arelocated on the control stand (P8)and the APU remote control panel(P62) on the nose landing gear.

Actuating either the APU fireswitch or external fire switch doesthe following:

• Closes the APU fuel valve.• Closes the APU bleed valve.• Trips the APU generator.• Arms the extinguishing system.• Silences the bell and resets the master warning lights.

Turning the APU fire handle (P8) ineither direction or pressing theAPU fire extinguisher bottle dis-charge switch on the lights/APU/interphone panel (P62) dischargesthe fire extinguisher bottle into theAPU compartment. Bottle dis-charge indication appears on theP8 panel.

The extinguisher bottle squib istested using the test switch on theright side panel (P61). Illuminationof the squib light indicates afunctioning squib.

APU Fire Extinguishing



Cargo Compartment FireDetection and Warning

Cargo compartment fire detectionis essentially accomplished with asmoke detection system that usesair sampling tubes and vacuumblowers to draw air through smokedetectors to activate warningsignals.

The forward cargo compartment isequipped with four 0.5 inch (1.3centimeter) diameter air-samplingtubes that extend from a main tuberunning along the left outboardside of the compartment. Twosmoke detectors connect to theforward end of the main tube andare attached to a vacuum chamberand blower assembly. With eithervacuum blower operating, air isdrawn from the cargo compartmentthrough the smoke detector andinto the chamber, where it isexhausted through the blower.

If smoke is present in the cargocompartment, it will be drawnthrough the sampling tube to thedetector and activate the alarm.

The aft cargo compartment isidentical to the forward in opera-tion with the, exception of addi-tional tubes to accommodate thelarger size of the aft compartment.The detectors are located in theright side of the compartment, justforward of the aft cargo door.

If both detectors in either compart-ment are activated, a fire alarm isinitiated. A fire in the cargo com-partment is indicated in the flightdeck by a fire bell; illumination ofthe master warning lights, discretefire warning light, and forward oraft cargo fire light; and a level Awarning on the EICAS display.

The complete cargo smoke detec-tion system can be tested beforeand during flight operation. Operat-ing the ENG/APU/CARGO testswitch activates the appropriatedetector and initiates the warn-ings.

Cargo Compartment Fire Detection and Warning



Cargo Compartment FireExtinguishing

There are two fire extinguisherbottles located in front of the aftcargo compartment. Each bottlehas two outlet fittings with explo-sive squib cartridges for releasingthe extinguishing agent.

Actuation of the appropriate cargofire switch, located on the controlstand (P8), arms the extinguishingsystem to direct the extinguishingagent into that compartment.Pressing the bottle dischargeswitch fires the selected squib andsends the agent into the compart-ment. The indicator light in theswitch illuminates when the bottleis discharged, which takes approxi-mately 30 minutes.

Pressing either cargo compart-ment-arming switch does thefollowing:

• Arms both bottle discharge switches.• Resets the fire bell and the master warning lights.• Turns off the forward or aft cargo heat fans.• Turns off one or both recirculat ing fans.• Opens the overboard exhaust valve (forward switch).• Turns off electric cargo com partment heater (aft switch).• Disables squib test function.

The extinguisher bottle squibs aretested via the SQUIB TEST switchon the test panel P61 A pressureswitch on each extinguisher bottleilluminates the DISCH light if thebottle is discharged or leaks belowa set pressure. The pressureswitch can be checked manually.

Cargo Compartment Fire Extinguishing



Wheelwell Fire and Duct LookDetection

The wheelwell fire warning isinitiated by a single sensor loopmounted on the ceiling of the mainwheel wells. Duct leak warning isinitiated by dual sensor loopsmounted along the pneumaticducts.

The alarm indications are initiatedwhen the wheelwell or duct tem-perature sensor reaches a prede-termined level. A wheelwell fireactivates the fire bell, the masterwarning lights, the discrete fire-light, and the discrete wheelwellfirelight. A duct leak turns on themaster caution lights and thediscrete duct leak lights on the P5pneumatic control panel andsounds the caution aural tone.

The EICAS provides caution andwarning messages on overheatand fire conditions. Because thereis no fire-extinguishing bottle forthe wheelwell, the landing gear islowered to put out a fire and coolparts such as brakes.

A duct leak can damage adjacent parts,so closing the appropriate valve shouldstop the airflow to that duct.

The wheelwell fire detection and duct leakoverheat detection systems can be testedfor continuity before and during flight.Operation of the DUCT LEAK test switchlocated on the right side panel (P61)causes the DUCT LEAK lights on theoverhead panel (P5) to illuminate if theseries detection elements and detectorcard circuitry are continuous.

Wheelwell Fire and Duct Look Detection

If the system is operating correctly,operation of the wheelwell test switch onthe fire/ overheat test module causes theWHL WELL FIRE light on the first officerspanel (P3-1) to illuminate. The discrete firelight and the master warning lightsilluminate and the fire bell sounds. TheEICAS also displays caution and warningmessages.



Freighter Fire Detection

Cargo compartment fire detectionis accomplished with a smokedetection system that uses airsampling tubes and vacuum blow-ers to draw air through smokedetectors to activate warningsignals.

The main deck cargo compartmentair is continuously sampled from18 pickup locations that flowthrough six smoke detector pairsto form two loops. If both loopssense smoke, the fire warningsystem is activated.

Because there are no extinguish-ers in the main deck or lowercargo area, a single switch is usedto depressurize all cargo compart-ments, which reduces the oxygenavailable for combustion. However,the flight deck continues to receivefresh, unpressurized air.

Main Deck Cargo Compartment Smoke Detection


Cabin Systems and LightingCabin Systems and LightingCabin Systems and LightingCabin Systems and LightingCabin Systems and Lighting

Features

EQUIPMENT AND FURNISHINGS

Cabin systems provide for thecomfor t and convenience of pas-sengers and crew members, forhandling and stowing cargo, andfor ensuring passenger and crewsafety in an emergency. Thesesystems include furnishings,lighting, oxygen, and lavatories.

WINDOWS

Fuselage windows are located inthree distinct areas on the 757: theflight deck; the passenger com-partment; and entry, service, andemergency doors.

DOORS

Access to various compartmentsand service areas is through entry,service, emergency, cargo, andaccess doors.

STOWAGE BINS

Each stowage bin supports amaximum of 180 pounds (82kilograms).

GALLEYS

The type of galley installed oneach airplane varies according tocustomer requirements. Galleysare usually customer-furnishedequipment.

EMERGENCY ESCAPE SYSTEM

The overwing exits are similar tothose used on the 707, 727, 737,and 767 airplanes.


Interior and exterior lighting ele-ments are almost identical on the757 and 767. The overhead stor-age bins for the two aircraft differonly in size. On the 757, the doorsopen on hinges in the usual way;on the 767, the doors slide into theceiling. Windows for the two mod-els are very similar.The 757 and 767 passenger ac-commodations and cargo systemsare essentially the same exceptthat the 767 have vacuum lavato-ries. Minor system differences,such as plumbing and compart-ment quantities, are required toaccommodate differences in fuse-lage size (single versus doubleaisles) and passenger counts.

• Overview

• Interior Arrangement

• Passenger Compartment

• Equipment and Furnishings

• Lighting

• Oxygen Systems

• Potable Water

• Lavatories and Galleys

• Doors

• Emergency Escape Systems



Overview

The cabin systems provide boththe necessities that make airplanetravel safe and the extras that helpmake it a pleasant experience.

Cabin systems include the cabinfurnishings, interior and exteriorlighting, oxygen, water and waste,and lavatory systems.

A description of the doors, emer-gency evacuation equipment, andfire protection equipment is in-cluded in this section.

The 757-200 interior uses indirectlighting and overhead stowagecompartments to obtain a spaciousand convenient arrangement.

The interior is adaptable to four-,five-, or six-abreast seating andcan accommodate a wide range ofseat pitches.

The overhead stowage bins pro-vide space for blankets, pillows,and most carry-on luggage. Theinterior volume of the 60-inch long(150 centimeter) bins is 9.63 cubicfeet (0.27 cubic meters). The binsreduce the amount of luggage thatmust be placed under the seats,providing a less cluttered interiorwith more foot space for the pas-sengers.

Passenger service units (PSU) arelocated under the stowage bins.Each unit has a reading light andan emergency oxygen mask foreach seat and an attendant callswitch.

For the freighter, all the passengerservice wiring, door indicatorcircuits, lighting, galleys, lavato-ries, PSU’s, and oxygen systemsare deleted.

Cabin Systems Overview



Interior Arrangement

In the basic airplane with fouroverwing emergency exits, seatsare arranged four abreast with acenter aisle in the first classsection and six abreast with acenter aisle in the economy sec-tion. Seating can range from 186passengers in a basic con-figuration to 224 passengers in ahigh-density configuration.

An option designated as thefour-door configuration replacesthe four overwing exits with twoemergency doors. The seatingranges from 178 to 239 passen-gers.

Seats are provided for five atten-dants throughout the cabin in thebasic configuration. High-densityconfigurations can have sevenattendant seats.

One lavatory is forward on the leftside for first class passenger use.The economy class lavatories arelocated in the aft cabin area. Allhave toilets, mirrors, washbasin,and other conveniences. Additionalor different locations for the lavato-ries may be selected.

The first class galley is locatedforward of the forward servicedoor, and the economy galleys arelocated forward of the aft servicedoor and across the full airplanewidth aft of the aft doors in a basicconfiguration. Additional or differ-ent locations for the galleys maybe selected.

Two closets are provided in thebasic airplanes. One is located aftof the forward service door andone aft of the No. 1 passengerdoor. A class divider is provided toseparate first and economyclasses. Additional or differentlocations for the closets and classdividers may be selected.

Typical Overwin Exit Configuratio Interior Arrangement



Passanger CompartmentEquipment and Furnishings

Sidewall panels are made ofNomex, Tedlar, and fiberglasssandwich construction. The win-dow reveal assemblies containsliding shades that are easilyremoved without removing thepanels.

Stowage bins are located alongboth sidewalls. The large doorsswing up to provide access. Emer-gency equipment is normally notstowed in the bins. Each bin sup-ports a maximum of 180 pounds(84 kilograms).

Above each seat is a PSU thatmay contain individual air outlets,oxygen masks and chemical oxy-gen generator, reading lights,attendant call button, speaker, andinformation signs. As an option,life vests or additional oxygengenerators and masks can bestowed in the PSU.

Passanger Compartment Equipment and Furnishings



Lighting

PASSENGER COMPARTMENTLIGHTING

General illumination is provided bydirect and indirect ceiling andsidewall fluorescent lighting and byincandescent wash lights adjacentto the lavatories and galleys. Low-level incandescent lights areprovided for night-lights. Entryarea lighting is provided by incan-descent threshold lights at eachentry and service door and byfluorescent lighting for the forwardleft entry door. General illumina-tion lighting is controlled at theforward left attendant station.Lighting adjacent to the flight deckdoor is dimmed when the door isopened to prevent glare. Thresholdlights are located at doors 1, 2,and 4.

Individual reading lights, located inthe PSU module, are provided foreach passenger. Incandescentwork lights at each attendant’sstation and fluorescent galley worklights provide specific area illumi-nation.

Incandescent dome lights that areilluminated when power is on inthe airplane, and fluorescentmirror lights that illuminate whenthe lavatory door is locked providelavatory lighting.

A call button is provided in eachlavatory to allow the occupant tonotify the attendants that assis-tance is needed.

NO SMOKING, FASTEN SEATBELT, and RETURN TO SEATsigns, which are visible at eachseat and lavatory, are controlledfrom the pilot’s overhead panel.The auto mode allows the NOSMOKING sign to illuminate whenthe gear is lowered and the FAS-TEN SEAT BELT and RETURN TOSEAT signs to illuminate when theflaps are lowered. A pressureswitch causes the NO SMOKINGand FASTEN SEAT BELT signs toautomatically illuminate at I 10,000foot cabin altitude.

Passander Comparment Lighting



SERVICE LIGHTING

Service lighting and associatedcontrols are located in the noseand main wheel wells; forward andaft cargo compartments, includingdoor and ceiling lights controllablefrom switches adjacent to eachdoor; air-conditioning bays; auxil-iary power unit (APU) compart-ment; tail cone compartment; mainequipment center; and groundfueling station.

Service Lights



EXTERIOR LIGHTING

Wing illumination lights aremounted on each side of thefuselage to light the leading edgeof the wing and the engine nacelle.

Four landing lights are installed:two in the wing roots and two onthe nose landing gear. The wingroot lights shine horizontally, andthe nose gear lights are aimeddownward on a typical glideslopeangle.

Two runway turnoff lights aremounted on the nose landing gearand illuminate the area to eitherside of the aircraft.

Optional taxi lights can be installedon either the fixed or movableportion of the nose landing gear.Strobe anticollision lights aremounted on the top and bottom ofthe fuselage and on each wingtip.The fuselage lights are coveredwith a red lens, and the wing lightsare covered with a clear lens.

Two position lights are mounted oneach wingtip facing forward andaft. The aft facing lights are cov-ered with a clear lens, and theforward facing lights have a redlens on the left wing and a greenlens on the right wing.

Four logo lights are installed in thehorizontal stabilizer and are posi-tioned to illuminate the verticalstabilizer and rudder.

Exterior Lighting



EMERGENCY LIGHTING

The emergency lights provideillumination of the passenger cabinand escape slide routes usingbattery packs, mounted in thecabin ceiling, as the power source.The lights can be switched manu-ally or set to come on automati-cally in the event of failure of thenormal airplane lighting power.

The lights consists of EXIT signmodules over each door, exitindicators near the floor at eachexit and over the main aisle be-tween doors, area lights on ceilingand floor in the cross aisles be-tween doors, main aisle lightsevenly spaced along the mainaisle, floor-mounted lights at20-inch (50-centimeter) intervalson the left side of the aisle, andslide lights externally mounted aftof each door and directed toilluminate the slide path.

Switches on the forward flightattendants or pilots panel controlall the emergency lights. Openinga door with the slide armed illumi-nates that exterior slide light.

Emergency Lighting



Oxygen Systems

The source of oxygen for the flightcrew is a single bottle of gaseousoxygen (see section 2, “FlightDeck”), and the sources for thepassengers and flight attendantsare multiple chemical generators.The systems are separate andindependent.

Passenger oxygen is provided bychemical oxygen generators in thePSU’s above the passenger andattendant seats and in lavatories.Oxygen mask deployment is initi-ated by an aneroid switch thatactuates when a cabin pressurealtitude of approximately 14,000feet or greater exists, or by manu-ally actuating a guarded switchinstalled on the pilots overheadpanel or manually opening eachmask stowage box.

Activation of the deployment circuitby the flight crew or by the aneroidswitch illuminates an indicator lighton the pilots overhead panel anddisplays a caution message on theengine indication and crew alertingsystem (EICAS).

Automatic or manual initiation ofmask deployment is sustainedthrough a timing circuit for 5 sec-onds to ensure release of all PSUdoors. Power application is re-moved after 5 seconds to eliminateunnecessary battery loading.

Pulling on any of the associatedoxygen masks activates oxygengeneration within each PSU. Eachchemical oxygen generator re-leases oxygen for a period of 12minutes (22 minute generators areavailable as an option).

Por table oxygen bottles are in-stalled throughout the passengercabin for therapeutic use.

Passanger Oxygen System



Potable Water

Potable water is supplied to thelavatory and galley areas from awater tank located aft of the aftcargo compartment. The basicwater tank contains 50 gallons(190 liters), but an optional 60-gal-lon (227-liter) tank is available. Thewater system is pressurized from acompressor receiving powerthrough a pressure switch. Thepneumatic system serves as abackup pressurization system.Heated drain masts drain thewater from the lavatory and galleyareas. Electrical heaters in thelavatory and galley water linesprovide hot water. Self-ventingfaucets are used to assist indraining the water system. Waterservice panels and a quantityindication system are incorpo-rated.

Portable Water System



Lavatories and Galleys

The basic configuration isequipped with four lavatories: oneadjacent to the forward passengerdoor and three in the aft cabin.Each lavatory contains a toilet,washbasin, mirror, and all thenecessary vanity items and dis-posal units. Other optional lavatorylocations are available.

The lavatories are equipped withaudible smoke detectors in theceiling and automatic fire extin-guishers and fire containmentsafeguards in the waste recep-tacles.

Ventilation is provided for eachlavatory through a blower locatedadjacent to the airplane outflowvalves.

The toilet in each lavatory includesa waste tank and flushing equip-ment. The tanks are servicedexternally through service panels.

The forward service panel is onthe bottom left side of the fuse-lage. Mid-cabin lavatories areserviced from a panel on the bodycenterline, a short distance aft ofthe wing-body intersection. The aftlavatories are serviced from apanel located on the centerline justaft of the main wheel wells.

Galleys are usually customerfurnished equipment and vary inaccordance with customer require-ments. A typical galley unit in-cludes complete food storage, hotmeal service, and coffee-makingequipment. The galley installationincludes privacy curtains to en-close the galley working area. Theforward galley area control panelincludes controls for both galleyand passenger cabin lightingsystems.

Cabin attendant panels, along withthe passenger address and flightinterphone systems, are located inthe vicinity of the cabin attendantseats.

Lavatory Waste System



Doors

GENERAL

The external door and hatcharrangement is shown. (Doors andhatches are plug type.) All doorsthat can be used as emergencyexits are operable from both insideand outside the airplane. Door sealinstallations are attached by me-chanical means for quick and easyreplacement. Each passenger andcargo door has a replaceablewear-resistant titanium thresholdscuff plate.

Each passenger door is equippedwith a pressurized gas poweredassist system for emergency dooropening.

Three plug-type passenger doorsare provided on each side of thepassenger compartment. Thedoors are manually operated andmove inward and then outwardwhen opened. These doors areclassified as Type I emergencyexits, in accordance with FAR 25.

For the overwing exit configurationtwo Type III inward-opening, re-movable overwing emergency exitsare provided on each side of thepassenger compartment. An op-tional arrangement installs Type Iemergency doors at STA 1335 inlieu of the overwing exits. Thesedoors are the outward-opening,fall-away type.

Each cargo compartment has anupward outward-opening door. Thedoors are opened and closed byelectrical drive units and latchedand unlatched by mechanicalmeans. Door controls are operablefrom either inside or outside thecargo compartment.

Access doors on the bottom of thefuselage allow entry to the lowersection 41, main electrical/elec-tronic center, aft body, and APU.

The door warning system, inconjunction with EICAS and thecaution and warning system, alertsthe flight crew whenever any dooris not closed and locked.

The freighter has a singleplug-type crew door on the leftside that swings inward on hinges.The outward opening cargo door ishydraulically operated and can beraised to a ver tical position.

Doors



ENTRY AND EMERGENCY DOORS

Six entry and two emergencydoors or six entry and four over-wing hatches are provided for thepassenger compartment. The threedoors on the left side are for entryand -exit and the three oppositeare normally for servicing. All areplug-type-hinged doors. Thesedoors feature a rotary snubber thatprovides snubbing on door openingor closing. Normal opening orclosing of the doors is assisted bysprings mounted on the fuselagetorque tube. An actuator and apressurized gas reservoir providepowered assistance for dooropening and slide deploymentupon door opening in the emer-gency mode. Opening from theoutside disarms the emergencyslide deployment and powereddoor opening assist system.The optional door aft of the wingson each side is for emergency useonly. It is hinged from the lowerdoorsill to open out and down, atwhich time the slide mounted on

the door deploys.

If overwing hatches are selected,the airplane is fitted with off wingescape slides that deploy auto-matically when a hatch is removedfrom inside the airplane.

Entry and Emergency Doors



Emergency Escape Systems

An escape slide is housed on eachexit door and covered by a con-toured lining (bustle). Except forthe optional No. 3 right and leftdoors, all slides will inflate whenthe door is opened in the armedemergency mode. No. 3 doors arearmed at all times and are usedonly for emergency exit. Thesedoors are plug types that open byfirst moving up, then out anddown, rotating around the bottomhinge, and finally hanging upsidedown.

The slide is inflated as the doorfalls. When the emergency armingmechanism is placed in the dis-armed position, or when the dooris opened from the outside, theslide is prevented from deploying.For over water operationsslide-rafts and rafts may be or-dered as options.

The overwing exits are inwardremovable hatches (located oneach side) that are similar to thoseused on the 707, 727, 737, and767 airplanes. An off wing escapeslide is stored in the wheelwellarea and is deployed automaticallywhen an overwing exit hatch isremoved from inside the passen-ger cabin.

The slide system is disarmed if thehatch is removed from the exterior.A secondary triggering control isavailable for deploying the slide.There is no need for a groundspoiler blow-down system sincethe slide and spoiler do not over-lap.

A variety of detachable emergencyequipment is located near exits inthe passenger cabin for easyaccess to handle different situa-tions.

Because the freighter has nopassenger doors or windows, theonly escape exits are through thecrew entry door (left side of flightdeck) and both No. 2 sliding win-dows in the flight deck, by meansof escape straps. External emer-gency operating handles areprovided on the crew entry doorand the right side sliding window.

Emergency Escape System, Four-Door Configuration



Typical Emergency Equipment Locations

Emergency Escape System, Overwing Exit Arrangement


Cargo SystemsCargo SystemsCargo SystemsCargo SystemsCargo Systems

Features

CARGO CAPACITY

The passenger airplane has twolower cargo compartments with acombined maximum weight capac-ity of 26,600 pounds (12 050kilograms) in 1, 790 cubic feet(50.7 cubic meters) of space.

CARGO DOORS

The main deck cargo door and theforward and aft lower lobe cargodoors are hinged at the top andopen outward. All cargo doors canbe operated manually if electric orhydraulic power is not available,and all are fitted with a continuouspressure seal to prevent the leak-age of pressurized air.

OPTIONAL TELESCOPING BULKCARGO SYSTEM

The bulk baggage loading systemwith telescoping modular bins isavailable as an option. The systemoffers easier loading and unloadingof the forward and aft cargo com-partments.


The cargo capacity and systems ofthe 757 and 767 differ significantlybecause of the large difference inpassenger capacity.

• Cargo Compartments

• Lower Lobe Cargo Doors

• Optional Telescoping Bulk Cargo System

• Main Deck Cargo Door

• Cargo Compartment Lights



Cargo Compartments

PASSENGER AIRPLANE

A total of 1,790 cubic feet (50.7cubic meters) is available in thetwo lower compartments: 700cubic feet (19.8 cubic meters) inthe forward compartment and1,090 cubic feet (30.9 cubicmeters) in the aft compartment.The weight limits are 10,300pounds (4650 kilograms) and16,300 pounds (7400 kilograms) inthe forward and aft compartments,respectively. Two 55-inch-wide(140 centimeter) outward-openingdoors are standard on the air-plane.

Both cargo compartments areclassified as Class C compart-ments, having both smoke detec-tion and fire extinguishing sys-tems.

An optional air cargo equipment(ACE) powered loading system isavailable for both cargo compart-ments. Reduced baggage andcargo damage and a reduction inramp personnel are advantages ofthis system. A powered belt load-ing system is also available.

FREIGHTER AIRPLANE

The freighter main cargo deck hasa volume of 6,600 cubic feet (187cubic meters) with a cargo doorsize of 134 by 86 inches (340 by218 centimeters). The main cargodoor control panel is located abovethe crew entry door and alsocontains the light switches for theentryway and cargo area. Manualoperation of the door is by a pumplocated in the left main wheelwell.

The lower cargo compartments arethe same as in the passengerairplane except that the aft com-partment is 40 cubic feet largerbecause of relocation of the aftelectronic center to the aft end ofthe compartment.

All cargo compartments are classi-fied as Class E, with smoke detec-tion systems but no fire extinguish-ing systems.

Cargo Compartments



Lower Lobe Cargo Doors

The forward and aft lower lobecargo doors are located on thelower right side of the fuselage.They are outward-opening doors,55 inches wide by 42.5 inches and44 inches high (140 centimeters by108 centimeters and 112 centime-ters), and weigh about 250 pounds(113 kilograms) each.

Cabin pressure loads on the doorsare carried by stop fittings on eachside of the door. A pressure seal isinstalled around the door to pre-vent air leakage when the door isclosed.

The doors are normally locked andunlocked manually and operatedelectrically from controls inside oroutside the cargo compartments. Amanual drive permits door opera-tion without electrical power.

Lower Lobe Cargo Doors



Optional Telescoping BulkCargo System

The bulk baggage loading systemsare located in the forward and aftcargo compartments of the air-plane. The systems use poweredtelescoping modules (bins) tofacilitate cargo loading and unload-ing.

There are three modules in the forwardcargo compartment and two in the aft.Module I in the forward compartment isthe largest of the three, and modules 11and III telescope into module 1. The threemodules in the forward compartment arelocated in the area aft of the forward cargodoor, and the two modules in the aftcompartment are located forward of theaft cargo door.

Optional Telescoping Bulk Cargo System



Main Deck Cargo Door

The main cargo door is on theupper left side of the forwardfuselage. It is hinged at the topand opens outward and upward toprovide a fuselage opening 86inches high by 134 inches wide(340 by 218 centimeters).

The door structure consists of apressure web forming the exteriorsurface, reinforced by frames. Acontinuous pressure seal aroundthe periphery of the door preventsoutward leakage of cabin pressureair and inward leakage of rain.Pressurization loads are transmit-ted to the fuselage structure bythe door hinge along the upperedge and by eight mechanicallatches along the lower edge.

The main cargo door is operatedwith hydraulic power from the lefthydraulic system. The door iscontrolled from the main cargodoor control panel above the crewentry door. The panel consists ofone guarded push-button switch

used to arm the door controlsystem and one momentary toggleswitch that controls door openingand closing. With electric andhydraulic power available and thedoor control system armed, mov-ing the door control switch toOPEN moves the door open untilthe switch is released or the doorreaches wide-open position. Withthe switch moved to CLOSE, thedoor closes until the switch isreleased or the door goes fullyclosed and latched. In addition,two EICAS messages are associ-ated with door operation.

A red DOOR OPEN light on thecontrol panel illuminates when thecargo door is unlocked and isextinguished when the door isclosed and locked. The controlpanel also contains cargo com-partment and entryway lightingand ground service switches.

The door opening or closing cyclecan be stopped and reversed atany point in the cycle. The door islocked closed through a hydrauli-

cally operated mechanical lock pinsystem.

The door can be operated in windsup to 40 knots and can stand inthe open position in winds up to 60knots.

The main cargo door can beopened manually if electric orhydraulic power is not available.The manual hydraulic hand pumpis a two-stroke pump used tomanually open and close the cargodoor. The door is opened by manu-ally positioning the door controlvalve lever to DOOR OPEN in theleft wheelwell and operating thehand pump until the door is raisedto the desired position. If batterypower is available, the DOOR NOTLOCKED indicator at the handpump illuminates when the door isunlocked. The door is closedmanually by placing the doorcontrol valve lever to DOORCLOSE and operating the handpump until the door is closed andlatched.Cargo Compartment Lights

Main Deck Cargo Door



Main deck cargo compartmentceiling lights are provided for theinterior of the main deck cargoarea. Threshold lights and exteriorlights are located adjacent to themain cargo door. Forward and aftcargo compartment lights areprovided for the interior and exte-rior of each compartment.

Cargo Compartment Lights


CommunicationsCommunicationsCommunicationsCommunicationsCommunications

Features

INTERPHONE SYSTEM

The communication systemsprovide communications betweenthe ground and the airplane. Com-munications are also providedbetween the crew and the passen-gers.

Cabin crew and ground crewcommunications are available aswell.

VOICE RECORDER

A voice recorder records the last30 minutes of flight crew communi-cations and conversations, to beretrieved in the event of an acci-dent.


Communication systems on the757 and 767 are vir tually thesame, with both using the sameline-replaceable units and commu-nication controls.

Both the 757 and 767 have dualVHF radios as basic equipment.The 757 and 767 add dual HFradios for the extended rangemodels. Single HF, dual HF, or HFprovisions in combination arefrequently ordered options. Cus-tomers ordering the ARINC com-munications addressing and re-porting system (ACARS) fre-quently order a third VHF radiodedicated to supporting theACARS equipment.

• Flight Interphone System

• Cabin/Service Interphone System

• Ground Crew Call System

• Passenger Address System

• VHF Communication System

• HF Communication System

• ACARS

• SELCAL System

• Voice Recorder System



Flight Interphone System

The flight interphone systemprovides a means of communica-tion between flight crew positions.The system also provides aninterface for the communicationand navigation systems via theaudio selector panels (ASP).

Hand microphones, boom micro-phones, or oxygen mask micro-phones can be connected throughthe ASP’s to the voice recorder,passenger address system, com-munication transceivers, and cabinand service interphone systems.The boom microphone or oxygenmask microphone is input to thevoice recorder if selected on theASP.

Jack outlets are provided at eachflight crewmember’s station formicrophones and headset. Pushesto talk (M) switches are located oneach ASP and on each controlwheel.

Flight Interphone System



Cabin/Service InterphoneSystem

The airplane communicationsystems provide internal communi-cations, as well as external com-munications to and from groundstations to facilitate ground servic-ing.

Interphone systems operate usingaudio selector panels, hand micro-phones or boom microphones,handsets, and service interphonejacks for communication betweenongoing maintenance functions.Passenger address (PA) systemvoice announcements are madeusing cabin handsets or audioselector panels. Nine serviceinterphone jacks are located at keyservice locations.

The cabin interphone providescommunications among the cabinattendants and between the cabinattendants and the flight crew.

This system also provides ameans of interfacing with theservice interphone and PA sys-tems.

Single-digit, touch-tone dialingfrom each handset and the pilotscall panel is provided to addresscalls to specific stations.

Handsets are located at all cabinattendant stations and the pilot’saft aisle stand (optional).

When a cabin location is dialedfrom one of the cabin interphonehandsets or the pilots call panel, apink call light is activated at thecalled station and a high/low chimeis activated through the warningelectronic unit (WEU).

When the flight deck is dialed froma cabin handset, a blue locationindicator light is activated on thepilots call panel and a high chimeis generated through the (WEU).When an alert call is made fromany handset, the pink lights at all

attendants’ stations flash continu-ously at 1-second intervals untilthe handsets are taken off thehook. At the same time, the high/low chimes are repeated threetimes through the PA amplifier. Thepink lights at each station ceaseflashing as the station’s handset istaken off the hook.

In the flight deck, an alert calllights a blue ALERT call light andsounds a high chime.

The handsets may communicatewith service interphone jacks, butflight deck personnel must selectthe SERV INTPH switch to ON forcommunication through the serviceinterphone jacks.

Cabin/Service Interphone System



The interior service jacks arelocated in three equipment areas.The exterior service jacks arelocated near the servicing areas ofthe airplane. The SERV INTPHswitch, located on the right sidepanel, must be switched on to hearmicrophone input signals from theservice interphone jacks. Theaudio selector panels providemicrophone input and headphoneoutput.

Ground Crew Call System

The ground crew call systemprovides both aural and visualsignals for use by the flight crew toalert the ground crew and for useby the ground crew to alert theflight crew.

A ground crew call switch is lo-cated on the pilots call panel.Operation of the ground crew callswitch sounds a horn in the nosewheelwell.

A pilots call switch is installed onthe nose gear panel. Operation ofthis switch sounds a single strokechime and illuminates a light in theflight deck.

Ground Crew Call System



Passenger Address System

The PA system provides flight crewcommunications to the passen-gers. It also provides audio tones(chimes) to alert attendants andpassengers.

PA inputs are prioritized. Flightcrew announcements have thehighest priority and are madeusing audio selector panels withany flight deck microphone. Atten-dants have priority 2 and use thecabin handsets to make announce-ments.

Priority 3 is prerecorded music orannouncements.

Priority 4 is boarding music fromthe boarding music/prerecordedannouncement tape reproducer.

Chimes are superimposed overany audio and are interpreted asfollows:

• Single low chime: No smoking, fasten seat belts.• Single high/low chime: Atten dant-to-attendant call.• Three high/low chimes: Atten dant alert.• Single high chime:

Passenger-to-attendant call.The PA audio output level is auto-matically increased when theengines are running or oxygenmasks are deployed to compen-sate for the increased cabin noise.

Passanger Address System



VHF Communication System

The very high frequency (VHF)communication system providesshort-range (line of sight) air-toground and air-to air voice anddata communications. Frequencyrange is 118 to 136 megahertz.

A dual VHF communication systemis basic. A third VHF transceiver isavailable as an option on basicmodels and is required on ex-tended range models.

Frequency selection for the VHFcommunication transceivers ismade from the respective VHFcontrol panel. Microphone selec-tion, headphone monitoring, andPTT functions are performed atthe audio control panels. Thecenter VHF communication system(optional) can be controlled by theARINC communications address-ing and reporting system (ACARS).When installed, ACARS uses theVHF system to receive and trans-mit digital data to and from aground station.

VHF Communication System



HF Communication System

The high-frequency (HF) communi-cation system is used forlong-range communications. SingleHF, dual HF or HF provisions arefrequently ordered options. EachHF communication system con-sists of a transceiver, a controlpanel, an antenna coupler, and acommon antenna.

The antenna and the antennacouplers are installed in the ver ti-cal fin leading edge. The antennacoupler matches the variableimpedance of the antenna to theimpedance of the transceiver overthe HF frequency range of 2.8 to30 megahertz in 1-kilohertz incre-ments.

Receiver sections of the HF trans-ceivers are protected for dual HFinstallations. When one HF trans-ceiver is transmitting, the other isprevented from receiving or trans-mitting.

System mode operation and fre-quency and sensitivity can beselected from the control panel.Audio connections to headset/speaker and microphone are madethrough the audio selector panels.Antenna tuning is automatic.

HF Communication System



ACARS (Optional)

ACARS provides a high-speeddigital data link between the air-plane and ground facilities. Bytransmitting and receiving dataautomatically without the flightcrew intervening, ACARS reducesflight crew workload. The system isused for exchange of airline opera-tions information such as flight andairplane identification; out of thegate, off the ground, on theground, into the gate reports(0001); delays, fuel, weather,airplane operating data, and so on.

ACARS is also capable of provid-ing voice telephone patch commu-nication between the airplane andground telephone circuits usingVHF radio, airline landlines,ARINC lines, and telephone sys-tems.

The main ACARS component isthe ACARS management unit(MU). The MU uses program pinsto determine airplane and airlineidentification and 0001 times. TheMU uses a VHF transceiver (nor-mally the center VHF transceiver)to receive and transmit data. Datafrom the VHF transceiver are toneencoded (1200 and 2400 hertz).The frequency used for datatransmission is 131.55 megahertz.Any frequency can be used forvoice communications.

The flight crew can provide somecontrol of the system by using aninteractive display unit (IDU).However, control can also comefrom a multipurpose control dis-play unit, (MCDU) or a dedicatedACARS control unit, depending onthe selected configuration. Themultipurpose printer is used by theflight crew to print ACARS repor tsstored in the MU.

A call light illuminates on the pilotscall panel and a single chime isgenerated by the WEU whenACARS receives a voice call. TheMU automatically tunes the VHFtransceiver to the voice frequencyselected by the ground station.When the call ends, the flight crewuses the IDU to revert ACARS tothe data mode. The flight crew canalso initiate a voice call to theground by using the IDU to providea frequency to the VHF trans-ceiver.

ACARS (Optional)



SELCAL System

The selective calling (SELCAL)system alerts the flight crew that aground station wishes to communi-cate with them. The alert is bymeans of a call light on the pilotscall panel and a single high chime.The light can be turned off and thesystem reset by pressing thecorresponding alert lamp/switch onthe pilots call panel or by keyingthe microphone PTT for that par-ticular receiver.

The SELCAL decoder acceptsinputs from the three VHF and twoHF communication systems.

The SELCAL decoder responds tothe ground station signal only ifthe signal is coded with theairplane’s unique SELCAL code.

SELCAL System



Voice Recorder System

The voice recorder provides acontinuous record of the last 30minutes of flight crew conversationand communication. It makes acontinuous 30-minute recording offour audio channels. The fourchannels are the captain’s, firstofficer’s, first observer’s, and thearea microphone on the voicerecorder control panel.

A bulk ERASE switch on the control panelcan be used on the ground with theparking brake engaged.

An underwater locator beacon is installedon the front of the voice recorder.

Voice Recorder System


Indicating and RIndicating and RIndicating and RIndicating and RIndicating and Recordingecordingecordingecordingecording

Features

FLIGHT DATA RECORDER

The flight data recorder systemrecords the last 25 hours of air-plane data (flight control surfaceposition, engine status, etc.) on amagnetic tape. These data helpdetermine the cause of an acci-dent.

WARNING AND ALERTING SYSTEM

The warning and alerting systemprovides the flight crew with visual,aural, and tactile indications ofabnormal airplane conditions.

ELECTRONIC CLOCKS

The clocks provide the flight crewwith a display of time, elapsedtime, and chronograph function.

• Digital Flight Data Recorder System

• Warning and Alerting System

• Electronic Clocks



Digital Flight Data RecorderSystem

The digital flight data recordersystem (DFDRS) provides thecapability to record the mostrecent 25 hours of flight param-eters on magnetic tape housed ina crash-proof container. The flightparameters include FAA manda-tory flight parameters and airlineoptional flight parameters.

FAA mandatory flight parametersinclude:

• Elapsed time• Pressure altitude• Computed airspeed• Vertical acceleration• Magnetic heading• Lateral acceleration• Pitch attitude• Roll attitude Horizontal stabi lizer position (pitch trim)• Thrust and power on each engine

• Trailing-edge flap positions (left, right)• Leading-edge devices

position (left, right)• VHF (radio transmitter) keying• Thrust reverser position• Speedbrake handle position• Marker beacon passage• Autopilot engagement• Longitudinal acceleration• Surface position primary con trols (pitch, roll, yaw)• Glideslope deviation• Localizer deviation• Autoflight control system mode and engagement status• Radio altitude• Master warning• Main gear squat switch status• Angle of attack• Outside air temperature or

Total air temperature• Hydraulics, each system, low pressure• Groundspeed

The required digital and analogdata from airplane flight systemsare supplied to the digital flightdata acquisition unit (DFDAU) forprocessing and format conversion.The digital output of the DFDAU isrecorded in the digital flight datarecorder (DFDR) on magnetictape.

An underwater locator beacon isinstalled on the front of the DFDR.A connector near the DFDR allowsconnection of a copy recorder foronboard readout of stored data.

Digital Flight Data Recorder



Warning and Alerting System

The warning and alerting systemprovides aural and visual warningand caution indications.

Airplane system signals are routedto the EICAS computer or thewarning electronic unit (WEU).When a non-normal condition issensed, appropriate visual andaural warnings are generated. Thesystem conditions that requireimmediate action are referred toas level “A” warnings. The systemconditions that require immediatecrew awareness and future actionare referred to as level B cautions.The system conditions that requireonly crew awareness are referredto as level C advisories.

The WEU is a card file. Electro-static discharge-sensitive (ESDS)device precautions should betaken to prevent damage to theelectronics on each WEU module.A wrist strap should always beused when removing and replacingWEU modules.

Warning and Alerting Systems



Electronic Clocks

Two clocks are installed in theflight deck, one each on thecaptain’s and first officer’s instru-ment panels.

Each clock provides readout ofGreenwich Mean Time (GMT) andprovides a GMT digital data signalor time reference to the flightrecorder and flight managementsystem.

Each clock also provides anelapsed time and a chronographfunction.

Electronic Clocks


NavigationNavigationNavigationNavigationNavigation

Features

ELECTRONIC FLIGHTINSTRUMENT SYSTEM

A primary flight display and navi-gation display for each pilotpresent all flight and navigationsystems information.

AIR DATA SYSTEM

Air data computers provide flightconditions to flight deck displaysand airplane systems.

WEATHER RADAR

A four-color presentation ofweather patterns that may includeturbulence is displayed on theelectronic horizontal situationindicator (EHSI).

GROUND PROXIMITY WARNINGSYSTEM

Terrain clearance and windshearannunciations warn the crew ofunsafe conditions.

INERTIAL REFERENCE SYSTEM

Ring laser gyro inertial referenceunits provide iner tial movementand attitude information.

FLIGHT MANAGEMENTCOMPUTER SYSTEM

Flight planning, position computa-tion, guidance, and performanceoptimization are supported by aworldwide database capability.


Navigation systems for the 757and 76 7 are nearly identical. Theflight management system andiner tial reference system are thesame.

Certified options available forthese systems allow configurationsto be tailored to each customer’sneeds. For example, acustomer-unique navigation data-base can be loaded into the flightmanagement computer. The data-base covers the geographic oper-ating area without altering thebasic functions and operations ofthe flight management computercontrol display unit, which is thepilots’ interface device.

• Flight Instrument System

• EFIS Displays

• Pitot-Static System

• Air Data Computer System

• Altitude Alert System

• Air Traffic Control System

• Traffic Alert and Collision Avoidance System

• Inertia] Reference System

• Weather Radar System

• Automatic Direction Finder System

• VHF Omnidirectional Range System

• Marker Beacon System

• Instrument Landing System

• Distance-Measuring Equipment

• Radio Altimeter System

• Ground Proximity Warning System

• Flight Management Computer System



Flight Instrument System

The flight instrument systemprovides airplane attitude, head-ing, vertical speed,distance-measuring equipment(DME) range, VHF Omnidirectionalrange (VOR) and automatic direc-tion finder (ADF) bearings, flightdirector and instrument landingsystem (ILS) commands, flightmanagement computer (FMC)displays, weather radar displays,radio altitude, and flight modeannunciations.

The electronic flight instrumentsystem (EFIS) uses cathode raytubes (CRT) for the primary flightinstruments. The electronic atti-tude director indicator (EADI) andthe electronic horizontal situationindicator (EHSI) are controlled byseparate symbol generators anddual control panels. Each pilot canindependently select an EHSIdisplay mode. The EFIS operateswith the flight management system(FMS) to provide the EADI’s withattitude and navigation information

The EHSI’s display ILS, VOR,MAP, or PLAN modes in a formsuitable for accurate and rapidreading by both pilots. The EFISalso provides visual indications offailure.

Selection of either the left orcenter symbol generator as a datasource for the captain’s EADI andEHSI is accomplished at thecaptain’s instrument source selectswitch. Similar control for the firstofficer’s EHSI and EADI using theright or center symbol generator isalso provided. With the otherswitches on the instrument sourceselect panels, both pilots canselect to alternate navigationsources in case of source failure.

The vertical speed indicators (VSI)receive inertial vertical speed fromthe inertial reference system(IRS).

The radio distance magneticindicator (RDMI) heading data aresupplied by the IRS, distance bythe DME systems, and bearingfrom either the VOR or ADF sys-tems.

Flight Instrument System



EFIS Displays

EADI MODES

The EADI has only one displaymode, in which airplane attitudeand flight director commands areshown. Speed tape (shown above)is an option.

EHSI MODES

The EHSI display depends on theposition of the mode select switchon the EFIS control panel.

The PLAN mode is generally usedbefore flight to review the route.The display is oriented north-up.The MAP mode is used duringflight to monitor airplane positionalong the route stored in memory.

The VOR modes are used whileflying a VOR radial. The ILS modesdisplay localizer and glideslopedeviations during approach andlanding. The two expanded modesdisplay only the horizontal situa-tion forward of the airplane. Thetwo FULL modes, which are op-tional on a basic airplane, displaya full compass rose. Weather datacan be displayed in EXP VOR,EXP ILS, and MAP modes.

EFIS Displays (Typical)



Pitot-Static System

The pitot-static system sensestotal and static air pressures andsupplies these pressures to theleft and right air data computers(ADC), the standby instruments,and other airplane systems asrequired.

The pitot probes and static por tssense both total air pressure andambient static air pressure. Thesepressures are used by the ADC todetermine the airplane’s altitude,airspeed, and other relatedparameters. The pressures arealso sent to the standby pneumaticairspeed and altitude indicators toprovide backup indications.

Pitot-Static System



Air Data Computer System

The ADC system monitors theenvironment around the airplane,processes the inputs, and sendsthe data to various airplane sys-tems. The system consists of twodigital ADC’s and the air datasensors. The air data sensorsinclude one total air temperature(TAT) probe, four pitot probes, twoangle-of attack (AOA) sensors,and two pairs of flush static ports.

Inputs to the ADC’s are total andambient air pressure from the pitotstatic system, barometric correc-tion from the altimeters, total airtemperature from the TAT probe,and angle of attack from the AOAsensors. The ADC’s convert theseanalog inputs to digital output airdata parameters.

The computer output signals aretransmitted on data buses thatsupply data to the air data instru-ments, engine and flight controls,navigation, warning, flight manage-ment, and Autoflight systems.

Air Data Computer System



Altitude Alert System

The altitude aler t system is in-stalled to alert the pilots to theapproach to or deviation from thealtitude selected on the autopilotmode control panel.

When the airplane approaches theselected altitude, the advisory lighton each electric altimeter lights.On deviation from the selectedaltitude, a level B warning is gen-erated, consisting of a level Bmessage on the engine indicationand crew aler ting system (EICAS),an aler t tone from the speakers,and illumination of the mastercaution lights and ALT ALERTcaution light. The altitude alertmodule receives barometric alti-tude from both ADC’s and is se-lectable with the captain’s ADCinstrument source select switch.

Altitude Alert System



Air Traffic Control System

The dual air traffic control (ATC)transponder system (left and right)contains the airborne componentsrequired to enable ground facilitiesto track airplane movementthrough the ground facility sectors.The ground facilities can monitorthe location, direction of travel,and altitude.

In response to interrogation pulsesreceived from a ground station, thetransponder replies with a pulsetrain providing identity information(selected code and identificationpulse if activated, mode A), andaltitude information (mode C)obtained from the ADC. The con-trol panel is used to set the as-signed code, select the left or rightsystem, and initiate an identifica-tion pulse. Each transponder isconnected to both ADC’s.

The selected transponder normallyuses its onside attitude sourceunless the onside ADC instrumentsource select switch switches it tothe alternate ADC.

The mode S address for the air-plane in which the transponder isinstalled is provided by the Dis-crete Addressable Beacon System(DABS) shorting receptacle.

The selective calling (mode S)transponder equipped airplanesand ground stations enhance theoperation of the ATC system byadding a discrete interrogationcapability and a data link feature,as well as performance improve-ments. The mode S transpondercan also function as part of anairborne separation assurance(ASA) system when interfaced witha traffic alert and collision avoid-ance system (TCAS).

Air Traffic Control System



Traffic Alert and CollisionAvoidance System

The TCAS is designed to alert theflight crew to the potential ofconflicts with other airplanes flyingin the same area. TCAS makesuse of existing ATC radar beaconsystems and the capabilities ofmode S transponders to coordi-nate with other TCAS equippedairplanes.

The system installed on currentcommercial airplanes is called TCAS 11.TCAS I provide two types of advisories tothe pilots. One type is the traffic advisory(TA), which is displayed on the EHSI andinforms the pilots that there are otherairplanes in the area.

The other type is the resolution advisory(RA), which is displayed on the EADI andadvises the pilots that a corrective orpreventive action is required to avoid anintruder airplane. TCAS 11 also providesaural alerts to the pilots.

The dedicated components of a TCAS 11system are a receiver/ transmitter, a topdirectional antenna, and a bottomantenna, which may be either directionalor omnidirectional.

TCAS II interfaces with the ATC systemand requires the use of mode Stransponders, top and bottom ATCantennas, and a mode S/TCAS controlpanel. TCAS 11 interfaces with the EFISand warning electronic unit (WEU) toprovide visual advisories and aural alerts.

Traffic Aler t and Collision Avoidance System



Inertial Reference System

The IRS is one of the airplane’sprimary sensing systems. As abasic sensor it provides airplanepitch and roll attitude, verticalspeed, body angular rates, andlinear acceleration. When it is inthe navigation mode it also pro-vides true and magnetic heading,present position, ground trackdirection and speed, and winddirection and speed.

The IRS consists of three inertialreference units (IRU) and aninertial reference mode panel(IRMP). The IRMP provides modeselection, control, and display;display of navigational parameters;and a means of initializing theIRU’s. The IRU’s sense angularrates around the pitch, roll, andyaw axes using ring laser gyrosand linear accelerations along thesame three axes using accelerom-eters.

Before the IRU’s can operate inthe navigation mode, they must gothrough a 1 10 minute alignmentperiod. During the alignmentperiod the airplane must not bemoved. The IRU’s must be initial-ized with the airplane’s accuratepresent position (latitude andlongitude) during this time. Initial-ization is normally done from theFMS control display units (CDU).Alternately, initialization can bedone using the IRMP.

Besides mode control and initial-ization, the only inputs requiredare air data parameters. The left orright ADC’s are selectable forinput.

IRS data are displayed on theelectronic flight instruments, theRDMI, and the VSI. Navigation,Autoflight, and other airplanesystems also use the data.

Inertial Reference System



Weather Radar System

The weather radar system pro-vides the pilots with an indicationof weather conditions along theirflight path, allowing them to diver ttheir flight around severe condi-tions. The pilots can also use theweather radar as a navigational aidby selecting the MAP mode anddisplaying prominent ground fea-tures such as coastlines, hilly ormountainous regions, cities, oreven large structures.

The transceiver generates radiofrequency pulses that are radiatedby the antenna.

The same antenna receives theradiated energy that is reflectedback by moisture-bearing clouds orby prominent terrain features. Thereflected radio frequency (RF)energy is processed by the trans-ceiver, displayed in color on theEHSI, and overlaid and scaled withthe navigation displays.

On-off control of the transceivers,as well as range selection andbrightness control, is accom-plished by the independent EFIScontrol panels for the captain andfirst officer. Different ranges canbe selected and displayed on thetwo display units.

The weather radar control panelincludes a mode selector (TEST,WX, WX+T (optional), and MAP),an antenna tilt control knob, and amanual gain control knob for usein MAP mode.

The antenna scan pattern is stabi-lized using attitude signals fromthe left IRU. When ALTN is se-lected on the captain’s IRS instru-ment source select switch, theattitude signals come from thecenter IRU.

Areas of light rainfall appeargreen, moderate rainfall appearsyellow, heavy rainfall appears red,and, as an option, the most turbu-lent areas of rainfall appear ma-genta.

A single system is shown above,but a dual system is optional.

Weather Radar System



Automatic Direction FinderSystem

The ADF system serves as anavigational aid to determinerelative bearing to a selectedground station with respect to theairplane centerline. It also providesaudio to the flight crew.

The ADF receiver can be tuned toreceive and compute a bearing toany radio transmitter with a fre-quency between 190.0 kilohertzand 1750.0 kilohertz. The receiversends the audio to the flight inter-phone system and the bearinginformation to the RDMI’s and tothe EFIS for display on the EHSI.

One ADF system is installed onbasic airplanes with two as anoption, and two ADF systems areinstalled on extended range air-planes.

Automatic Direction Finder System



VHF Omnidirectional RangeSystem

The airplane’s VOR system pro-vides navigation data for the flightcrew and the FMS. The VORsystem consists of two VOR con-trol panels and two VOR navigationreceivers. The flight crew selects aVOR navigation ground station bytuning in a VOR frequency and thedesired course to the VOR stationon the VOR control panel.

The VOR station constantly trans-mits magnetic bearing informationin an omnidirectional (all direc-tions) pattern. The VOR receiver inthe airplane receives and decodesthe signal to determine the mag-netic bearing of the airplane to theVOR station.The VOR receivers send magneticbearing data to RDMI’s and to theairplane EFIS for display. On theRDMI’s, a needle points to themagnetic bearing of the VORstation. VOR deviation is displayedon the EFIS EHSI’s. VOR deviation

is the difference between thecourse that has been selected onthe control panel and theairplane’s magnetic bearing to theVOR station.

During the ILS and VOR EFISmodes, the VOR is manually tunedand an indicator light on the con-trol panel shows MAN. When MAPor PLAN is selected on the EFIScontrol panel, the indicator lightshows AUTO and the FMC auto-matically tunes (auto tunes) theVOR receivers. The FMC’s can usethe bearing data from the VORreceivers in conjunction with thedistance information from a DMEinterrogator for position updating inflight.

VHF Omnidirectional Range System



Marker Beacon System

The marker beacon system indi-cates to the flight crew that theairplane is passing over a particu-lar geographical location such as apoint along an instrument landingpath. The marker beacon receiveris a module within the VOR re-ceiver that is enabled only in theleft system.

Marker beacon transmitters lo-cated in standard flight pathstransmit a narrow vertical beam oftone-modulated 75 Megahertz RFAs the airplane flies over a beam,the receiver detects the RIF andlights the appropriate panel light.During runway approaches, threetypes of markers can be used withpanel lights illuminating and tonessupplied to the pilots.

Marker Beacon System



Instrument Landing System

The ILS provides precision ap-proach guidance during instrumentapproaches by giving position datarelative to the glidepath and run-way centerline.

Three complete systems areinstalled, left, center, and right, allcontrolled by a single control panelused for frequency and frontcourse selection.

Position-sensitive radio signals arereceived from the glideslope andlocalizer transmitters. The ILSreceiver, which computes devia-tion, either up or down, right or leftdecodes the signals. Deviationdata are sent in digital format tothe FMC, autopilot, and groundproximity warning system (GPWS).Deviation is also sent to the EFISand displayed on the EADI andEHSI displays. The center systemsends data to the standby ILSdisplay. Audio is sent to the flightinterphone system for stationidentification.

Instrument Landing System



Distance-Measuring Equipment

A dual (left and right) DME systemis installed to provide the pilotswith the slant range distancebetween the airplane and a groundstation. The distance is displayedin digital format on both the RDMIand EHSI. Continuous distanceinformation is also provided to theFMC for position updating in flight.

The DME interrogator transmits apulse pair to a selected groundstation, which then retransmits thepulse pair to the airplane. The timetaken for the round-trip signal ismeasured, and distance is thencomputed. The DME station isselected using the VOR controlpanel in VOR mode and the ILScontrol panel in the ILS mode formanual tuning. In the MAP orPLAN modes the FMC automati-cally tunes both DME’s. DMEfrequencies are paired with theVOR and ILS frequencies.

Distance-Measuring Equipment



Radio Altimeter System

The radio altimeter system pro-vides the pilots with an indicationof the terrain clearance altitude.The system is used at low altitude(0 to 2,500 feet), primarily duringapproach, landing, and takeoff.

The system consists of threereceiver/transmitter units, eachwith its own transmit antenna andreceive antenna. The receiver/transmitter unit computes thealtitude, which is then displayed onthe EADI’s.

Each pilot can select a decisionheight (DH) altitude from theonside EFIS control panel fordisplay on the EADI above theradio altitude display. When theradio altitude is equal to or lessthan the DH, the DH displaychanges color and size and mo-mentarily flashes DH. Radio alti-tude is used by the autopilot flightdirector system, GPWS, warningsystem, and EICAS.

Radio Altimeter System



Ground Proximity WarningSystem

The GPWS provides visual andaural alerts whenever the airplaneis in danger of contacting theterrain. When the airplane is below2,500-foot radio altitude the GPWScontinuously monitors terrainclearance altitude, descent rate,terrain closure rates, glideslopedeviation, and flap and gear con-figuration so that caution andwarning alerts can be generated ifthe airplane is unsafe because ofcloseness of terrain.

The GPWS includes a groundproximity warning computer(GPWC), flap and gear overridepanel switches, a flight deck testswitch, and WINDSHEAR, PULLUP, and GND PROX panel lights.The GPWC processes inputs fromthe radio altimeter, air data, iner-tial reference and instrumentlanding systems, as well as flap,slat, and gear configuration, toprovide alerts in GPWS modes 1

through 6. It receives additionalinput from the stall warning com-puters for the generation of mode7 (windshear) warning.

Flap and gear override switchesallow the pilots to inhibit or cancelaler ts that are generated while theflaps or landing gear are intention-ally out of normal configuration.The GND PROX light is also aswitch that can be pushed toinhibit or cancel the mode 5 glide-slope aler t.

The GPWC sends outputs, de-pending on the GPWC mode, to itspanel lights and to the WEU, whichcontrols the master warning lights.It sends a signal to the EFIS forwindshear warnings, which displaythe red word WINDSHEAR on theEADI. The aural alerts, which aregenerated inside the GPWC, arealso sent to the WEU for amplifica-tion and fur ther routing to the auralwarning speakers.

Ground Proximity Warning System



Flight Management ComputerSystem

The flight management computersystem (FMCS) consists of twoFMC’s and two CDU’s. The FMCSuses inputs from navigation,engine, fuel system, and flightenvironment sensors along withstored and entered data to performflight crew selectable operations.

Flight Management Computer System

The flight crew can enter, modify,or retrieve data and select modesfrom the CDU. The FMC has thecapability to store any desiredlateral and vertical flight plan andto compute guidance, navigation,and performance commandsrelative to the flight plan. The FMCcan be used to provide advisorydata to the flight crew, enablingthem to fly a selected course orprofile. The FMC can also becoupled to the automatic flightcontrol systems to automaticallyfollow a planned flight profile.

The FMCS provides a map displayand selected navigation informa-tion for display on EFIS displays.The FMCS also provides thrusttarget cursor control to the EICAS.


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Features

AUTOPILOT FLIGHT DIRECTORSYSTEM

The autopilot flight director systemProvides flight director commandsor automatic control of the aircraftin all phases of flight. It generatesstabilizer trim commands when theautopilot command is engaged.

THRUST MANAGEMENT SYSTEM

The thrust management systemprovides full flight regime controlof throttles, including takeoff.

YAWDAMPERSYSTEM

The yaw damper system elimi-nates yaw rates associated withDutch roll.

STABILIZER TRIM SYSTEM

The stabilizer trim system providesspeed or Mach trim control whenthe autopilot is not engaged. Itaccepts autopilot trim requestsduring command engagement.

MAINTENANCE MONITOR SYSTEM

The maintenance monitor systemprovides a single point for deter-mining automatic flight controlsystem faults.


Autoflight systems on the 757 and767 are fundamentally identical;both use the same flight manage-ment computer with softwarespecifically tailored for each con-figuration. Both airplanes featurethe same Cat III b autoland sys-tem as basic equipment. Thisadvanced function allows landingin low-visibility conditions (150 feet(50 meters)) with zero decisionheight.

• Autopilot Flight Director System

• Thrust Management System

• Yaw Damper System

• Stabilizer Trim System

• Maintenance Monitor System



Autopilot Flight Director System

The autopilot flight director system(AFDS) provides flightcrew-selectable modes for flightpath control. System operation iscontrolled from the mode controlpanel, with armed and operatingmodes displayed on the electronicairborne data indicators (EADI).

Three flight control computers(FCC) are installed, each control-ling dedicated pitch, roll, anddirectional control servos (autopi-lot controls the yaw directionalcontrol servos only during multi-channel approach). All functionsfor three-axis control of the aircraftare contained in each FCC. Eachcomputer provides automaticstabilizer trim commands to astabilizer trim and elevator asym-metry module (SAM).

Autopilot operating modes areselected on the Autoflight controlsystem (AFCS) mode controlpanel. The FCC’s use selectedmode, navigation sensor inputs,and flight management computer(FMC) inputs to generate outputsignals.

These signals can be used forflight director display only, or maybe used to actively control theaircraft.

Flight director commands aredisplayed on the EADI’s. Theactive AFDS mode is also dis-played on the EADI’s.

Autoland status annunciatorsprovide indications of autopilotsystem status including degrada-tion from Cat III landing capability.

Autopilot disengagement gener-ates a visual and aural annuncia-tion (EICAS level A warning).Reduced system capability resultsin an autopilot caution (EICASlevel B caution). The maintenancecontrol and display panels (MCDP)record all flight faults for post flightmaintenance review and can beused to ground-test the flightcontrol computers and interfacingsystems.

Autopilot Flight Director System



Thrust Management System

The thrust management systemperforms thrust limit calculationsand auto throttle functions. Itcontrols the throttles through thefull flight regime of operation. Onethrust management computer(TMC) is installed in the aircraft toperform both operations,

Thrust limits are establishedmanually from the thrust modeselect panel or automatically onpower-up from the FMC in VNAVmode. The AFCS mode controlpanel is used to select the appro-priate auto throttle modes. Theauto throttle functions depend onthe mode selected and can controlthrust, Mach, airspeed, rate ofaltitude change, or throttle retardrate. Thrust limit protection isactive in all modes to preventoverboost, overspeed, or minimumspeed exceedances.

Auto throttle modes are displayedon the EADI’s. Thrust limit modesare displayed on EICAS upperdisplay units. Disengagement ofthe auto throttle causes visual andaural indications. The systemsends flight fault information to theMCDP for post flight analysis.

Thrust Management System



Yaw Damper System

The yaw damper system generatesrudder commands to dampenundesired yaw based on input fromair data computers and the inertialreference system. Modal suppres-sion accelerometers provide anadditional input that is used toimprove passenger comfort.

The yaw damper control panelprovides on off control and systemstatus indications. Inoperativestatus is also annunciated on theEICAS upper display unit with alevel C message L YAW DAMPERor R YAW DAMPER indicating theinoperative system.

The yaw damper modules use datainput from air data computers,iner tial reference units, and modalsuppression accelerometers tocompute rudder commands appro-priate to existing flight conditions.

These commands go to the yawdamper servos. The module alsomonitors system operation andperforms both manually initiatedand automatic system testing.

A 12-character LED display oneach yaw damper module is usedto display preflight test results andexisting and last flight leg yawdamper system faults.

The yaw damper servos use elec-trical commands from the yawdamper modules to control hydrau-lic flow to an actuator piston. Thismechanical servo output issummed with any manual or auto-pilot rudder commands. Maximumrudder authority is 3 degrees leftor right for each yaw damper. Theoutputs of the left and right sys-tems are mechanically summed fora total deflection of ±6 degreeswith both systems active. The yawdampers are independently testedusing a yaw damper test switchlocated on the test panel. Testingthe yaw damper causes the rudderto move ±3 degrees.

Yaw Damper System



Stabilizer Trim System

Aircraft longitudinal trim control ismaintained by the stabilizer trimsystem. The horizontal stabilizer isa movable airfoil that is normallycontrolled by the SAM. As theairplane center of gravity changesduring flight, the angle of attack ofthe stabilizer is varied to controlairplane longitudinal attitude.

Airplane pitch axis trim is main-tained during flight by several trimmodes.

In the alternate electric manualtrim and manual electric trimmodes the flight crew controlstrim.

In the speed trim mode the stabi-lizer is automatically trimmed atlower airspeeds when flaps are notretracted.

In the Mach trim mode the stabi-lizer is automatically trimmed asMach increases when the flaps areretracted. Mach input is receivedfrom the air data computer.

In the auto trim mode the stabilizeris trimmed by commands from theengaged FCC.

Stabilizer position indicators onthe aisle stand provide visual trimindications to the flight crew.

Automatic stabilizer trim, Machtrim, speed trim, and manual trimmodes all use common systemcomponents. The electric alternatemanual trim is sent directly to thestabilizer trim control modules. Allother trim commands are sentthrough the SAM. Hydraulic poweris used to move the stabilizer, andcutout switches allow systemshutdown in the event ofuncommanded stabilizer motion.

Stabilizer Trim System



Maintenance monitor System

The maintenance monitor systemcombines flight fault storage andground test functions for the AFDSand thrust management systems.The system also provides flightfault storage for the FMC.

One maintenance control anddisplay panel (MCDP) is installed.It is connected to the FCC’s,TMC’s, and FMC’s. Flight faultdata can be displayed for all sixcomputers. The MCDP is also usedto perform ground testing of theFCC’s and TMC. Avionics systemsthat interface with these comput-ers are also checked during inter-face tests.

The MCDP is normally off. It turnson automatically at the end of aflight, and all the connected com-puters transmit messages detailingany failures that occurred duringthe flight. The MCDP automaticallyshuts down after all flight faultinformation has been stored or 3minutes have elapsed. Mainte-nance personnel can use theMCDP to review flight faults andalso to perform ground tests forfault isolation or system verifica-tion.

The MCDP can be operated fromthe main equipment center usingfront panel controls. Wiring provi-sions are installed to connect aremote control panel in the flightdeck to the MCDP for control, withall messages displayed on theEICAS lower display.

Maintenance monitor System


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AABSV alternate brake selector valveac alternating currentACARS ARINC communications

addressing and reportingsystem

ACC active clearance controlACE air cargo equipmentACM air cycle machineACMP alternating current motorADC air data computerADF automatic direction finderADI attitude direction indicatorAFCS autoflight control systemAFDS autopilot flight director systemAIV accumulator isolation valveAPU auxiliary power unitARINC Aeronautical Radio, Inc.ASAS airborne separation assurance

systemASP audio selector panelATC air traffic controlAVM airborne vibration monitor

BBITE built-in test equipmentBL buttock lineBMV brake metering valveBPCU bus power control unitBTB bus tie breakerBVCU bleed valve control unit

CC centerCAA Civil Aviation Authority (UK)CDU control display unitCRT cathode ray tubeCSEU control system electronic

DDABS discrete addressable beacon

systemDADC digital air data computerdB decibelDC direct currentdeg degreeDFDAU digital flight dataDFDR digital flight data recorderDFDRS digital flight data recorder

systemDH decision heightDME distance measuring ringDSP display select panel

EEADI electronic attitude directorECU electronic control unitEDP engine-driven pumpE/E electrical/electronicEEC electronic engine controlEFIS electronic flight instrument

systemEGT exhaust gas temperatureEHSI electronic horizontal situation

indicatorEHSV electrohydraulic servo valveEICAS engine indication and

crew alerting systemEPCS electronic propulsion control

systemEPR engine pressure ratioESDS electrostatic discharge sensitiveETOPS extended-range twin operations

FFAA Federal Aviation AdministrationFADEC full-authority digital electronic

controlFAR Federal Aviation RegulationFCC flight control computerFFG fuel flow governorFIS flight instrument systemFMC flight management computerFMCS flight management computer

systemFMS flight management systemFQIS fuel quantity indicating systemFSEU flap/slat electronic unitft footFwd forward

Ggal gallonGCB generator circuit breakerGCR generator control relayGCU generator control unitGMT Greenwich mean timeGPWC ground proximity warning

computerGPWS ground proximity warning

system

HHF high FrequencyHMG hydraulic motor generatorHOT high oil temperatureHP high pressureHPC high pressure controllerHPT high pressure turbineh hourHSI horizontal situation indicatorHz hertz

IIDG integrated drive generatorIDU interactive display unitIGV inlet guide vaneIGVA inlet guide vane actuatorILS instrument landing systemin inchinbd inboardIP intermediate pressureIRMP inertial reference mode panelIRS inertial reference systemIRU inertial reference unit

Kkeas knots equivalent airspeedkHz kilohertzkt, kn knotkVA kilovoltampere

LL left, literlb poundLCD liquid crystal displayLCIT load compressor inlet

temperatureLE leading edgeLED light-emitting diodeLOP low oil pressureLP low pressureLPT low-pressure turbineLRU line-replaceable unitLVDT linear variable differential


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MMCDP maintenance control andMCDU multipurpose controlMCP mode control panelMEC main equipment centerMES main engine startmHz megahertzmin minutemph miles per hourMU management unit

NN

1, N

2 , N

3 rotor assembly number

NM, nmi nautical mileNVM nonvolatile memory

OOOOI out of the gate, of the ground

operations, on the ground, intothe gate

Outbd outboard

PP&W Pratt & WhitneyPA passenger addressPCA power control actuatorPCU passenger control unit, power

control unitPDIU propulsion discrete interface

unitPDU power drive unitPMA permanent magnet alternatorPRSOV power regulating and shut-off

valvePSEU proximity switch electronic unitpsi pounds per square inchpsig pounds per square inch gaugePSM power supply modulePSU passenger service unitPSUD passenger service unit displayPTT push to talkPTU power transfer unit

RR rightRA resolution advisoryRAT ram air turbineRBS radar beacon systemRCM ratio changer moduleRDMI radio distance magnetic indica

torRF radio frequencyrpm revolutions per minuteR-R Rolls-RoyceRVDT rotary variable differential

SSAM stabilizer trim and elevator

assymetry moduleSCM spoiler control moduleSEB seat electronics boxsec secondSEI standby engine indicatorSELCAL selective callingSG symbol generatorSIL speech interference levelSPM stabilizer position moduleSTA stationSTCM stabilizer trim control moduleSWC stall warning computer

TTA traffic advisoryTAI thermal anti-iceTAT total air temperatureTCAS traffic alert and collisionTLA thrust lever angleTMC thrust managementTRU transformer/rectifier unittyp typical

VV voltVHF very high frequencyVOR VHF omnidirectionalVSI vertical situation indicatorVSV variable stator vane

WW waftWBL wing buttock lineWEU warning electronic unitWL water lineWS wing stationWXR weather radar

YYDM yaw damper module


B-757-General-Familiarization for Maintenance and Pilots

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