Embraer Emb-135bj Aom Vol. 2 Aom

Copyright 2001 by EMBRAER - Empresa Brasileira de Aeronáutica S.A.. All rights reserved. This document shall notbe copied or reproduced, whether in whole or in part, in any form or by any means without the express written

authorization of Embraer. The information, technical data, designs and drawings disclosed in this document areproperty information of Embraer or third parties and shall not be used or disclosed to any third party without

permission of Embraer.

AIRPLANE OPERATIONS MANUALVOLUME 2

EMPRESA BRASILEIRA DE AERONÁUTICA S.A.

This manual is applicable to the EMB-135BJ airplanes allmodels equipped with AE3007A1E engines, operating underFAA certification.

NOTE: THE EMB-135BJ MODEL HAS THE COMMERCIALDESIGNATION OF LEGACY.

THIS PUBLICATION IS INCOMPLETE WITHOUT AIRPLANEOPERATIONS MANUAL AOM–135/1542 VOLUME 1

AOM–135/1542-07NOVEMBER 30, 2001

REVISION 7 – MARCH 31, 2005

TRANSMITTAL LETTER

Subject: REVISION 7 of the AOM–135/1542-07 Airplane OperationsManual

Dear Madam or Sir,

You are receiving REVISION 7 of the AOM. To help you to betterunderstand this revision, we are presenting the main points in the“Highlights of Change” inside the manual.

This revision incorporates all preceding temporary revisions. Thereforeall yellow pages should be discarded from the manual.

The revised pages supersede the current ones. Therefore, the affectedpages must be replaced and the new pages inserted following thenumbered sequence.

The Section 1-03-36 code 01 is being deleted. Current pages must beremoved from manual and discarded. The information must bereplaced by Section 1-03-36 code 02.

In case of any question, please contactEmbraer Flight Operations Engineering Department byE-mail: [email protected] or by phone: (+55 12) 3927 1706

THIS PAGE IS NOT PART OF THE MANUAL AND MUST NOT BEINCORPORATED IN YOUR AOM. PLEASE DISCARD IT AFTERREADING.

CUSTOMER COMMENT FORMAirplane Operations Manual

AOM–135/1542

Please use this Customer Comment Form to notify us of

any discrepancies or problems you find in the Airplane

Operations Manual. We would also welcome constructive

suggestions on how we can further improve our

documentation or service.

Your feedback will be acknowledged, and we will advise

you of the action we intend to take.

Sincerely,

Embraer Operations Support

Please return this form to:

Embraer - DSF/GSO/SEO - PC176Av. Brigadeiro Faria Lima, 2170

CEP 12227-901São José dos Campos - SP - BRASIL

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Phone: +55 12 3927-1706Fax: +55 12 3927-2477E-mail: [email protected]

Airplane Operations ManualAOM–135/1542

Section/Page/Code: Date of Page: Revision:

Name:Position:Company:Fax Number: Phone Number:E-Mail:

Comment/Suggestion:

Space reserved for Embraer CCF nº:

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SECTION 2-13

FLIGHT CONTROLS

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-13-05 ..01Pitch Control....................................................................... 2-13-10 ..01

General........................................................................... 2-13-10 ..01Elevator .......................................................................... 2-13-10 ..02

General ....................................................................... 2-13-10 ..02Jammed Elevator........................................................ 2-13-10 ..02Jammed Elevator Operation ....................................... 2-13-10 ..02

Tabs................................................................................ 2-13-10 ..02General ....................................................................... 2-13-10 ..02Servo Tabs ................................................................. 2-13-10 ..02Spring Tabs ................................................................ 2-13-10 ..02

Pitch Trim System .......................................................... 2-13-10 ..04General ....................................................................... 2-13-10 ..04System Components .................................................. 2-13-10 ..04

Horizontal Stabilizer Control Unit (HSCU)............. 2-13-10 ..04Horizontal Stabilizer Actuator (HSA) ..................... 2-13-10 ..04

System Operation ....................................................... 2-13-10 ..04Pitch Trim Channels Priority ....................................... 2-13-10 ..06Pitch Trim System Protection ..................................... 2-13-10 ..06

Switch Protection................................................... 2-13-10 ..06Runaway Protection .............................................. 2-13-10 ..06Inadvertent Actuation Protection ........................... 2-13-10 ..06HSA Excessive Load Protection............................ 2-13-10 ..07

EICAS Messages ........................................................... 2-13-10 ..08Controls and Indicators................................................... 2-13-10 ..10

Control Stand.............................................................. 2-13-10 ..10Control Wheel ............................................................. 2-13-10 ..11Control Pedestal Aft Panel.......................................... 2-13-10 ..12EICAS Indication......................................................... 2-13-10 ..14

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Roll Control.........................................................................2-13-15.. 01Aileron Control System ...................................................2-13-15.. 02Roll Trim System ............................................................2-13-15.. 04EICAS Messages............................................................2-13-15.. 06Controls and Indicators...................................................2-13-15.. 06

Flight Controls Panel ..................................................2-13-15.. 06Control Stand..............................................................2-13-15.. 08Control Pedestal Aft Panel..........................................2-13-15.. 10EICAS Indications.......................................................2-13-15.. 11

Yaw Control ........................................................................2-13-20.. 01Rudder Control System ..................................................2-13-20.. 02

Automatic Shutoff Through the Speed Switch............2-13-20.. 04Rudder Hardover Protection.......................................2-13-20.. 04

Rudder Deflection ...........................................................2-13-20.. 05Airplanes Under CTA and FAA Certification...............2-13-20.. 05Airplanes Under JAA Certification ..............................2-13-20.. 05

Yaw Trim System............................................................2-13-20.. 06EICAS Messages............................................................2-13-20.. 08Controls and Indicators...................................................2-13-20.. 09

Flight Controls Panel ..................................................2-13-20.. 09Control Pedestal Aft Panel..........................................2-13-20.. 10Main Panel..................................................................2-13-20.. 11EICAS Indications.......................................................2-13-20.. 12

Gust Lock System ..............................................................2-13-25.. 01Electromechanical Gust Lock System ............................2-13-25.. 01

Locking Operation ......................................................2-13-25.. 02Unlocking Operation ...................................................2-13-25.. 04

Controls and Indicators...................................................2-13-25.. 06Glareshield Panel .......................................................2-13-25.. 06Control Stand..............................................................2-13-25.. 07

Flap System........................................................................2-13-30.. 01Flap System Operation ...................................................2-13-30.. 02EICAS Messages............................................................2-13-30.. 04Controls and Indicators...................................................2-13-30.. 04

Control Pedestal Aft Panel..........................................2-13-30.. 04EICAS Indications.......................................................2-13-30.. 06

Spoiler System ...................................................................2-13-35.. 01Ground Spoiler................................................................2-13-35.. 02Speed Brake ...................................................................2-13-35.. 02EICAS Messages............................................................2-13-35.. 04Controls and Indicators...................................................2-13-35.. 04

Control Stand..............................................................2-13-35.. 04EICAS Indications.......................................................2-13-35.. 06

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GENERALThe primary flight control system consists of elevators, ailerons andrudder. Elevators are mechanically actuated. The ailerons and rudderare hydraulically powered and may also be mechanically actuated incase of loss of both hydraulic systems.

Trim system is provided in all axis. Tabs are provided for pitch controlonly, and are not available for ailerons and rudder.

A gust lock system blocks elevator controls on the ground, avoidingdamage to the control systems in case of strong wind gusts. Therudder and ailerons are hydraulically damped for the same purpose.

An electrically operated flap system is provided with five discretepositions.

Speed brakes installed overwing allow increased descent rate and helpin decelerating the airplane. Ground spoilers destroy lift, thus providingbetter braking effectiveness.


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PITCH CONTROL

GENERALPitch control is provided by mechanically-actuated elevators and anelectrically-positioned horizontal stabilizer which is commandedthrough the Pitch Trim System. Tabs are automatically positioned, thusreducing pilots effort.


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ELEVATOR

GENERAL

The primary pitch control system is performed by the elevators, whichare actuated through a fully duplicated set of command circuits.

JAMMED ELEVATOR

In case of jamming of one of the circuits (left or right), both elevatorpanels may be disconnected through a handle located on the controlpedestal. This procedure will release the free elevator panel from itsjammed counterpart, allowing the free panel to be commanded.When disconnected, an amber light illuminates on the control stand.Controls cannot be reconnected during flight, requiring maintenanceaction.

JAMMED ELEVATOR OPERATION

The autopilot elevator servo and the stick pusher servo are connectedon the left side of the disconnection device. Once disconnection isactuated, the stick pusher will actuate only on the left side and autopilotmust not be used.

TABSGENERAL

There are four tabs, two on each elevator panel, located near theelevator root. The outer tabs are servo tabs and the inner tabs arespring tabs.

SERVO TABS

The deflection of the servo tabs is proportional to the elevatordeflection. Since the servo tabs proportionally deflects in the oppositedirection to the elevators, it promotes a reduction in the forcesrequired.

SPRING TABS

The spring tabs are connected in such a way that elevator deflection inone direction causes the spring tab to move in the opposite direction,thus reducing the amount of force required to move the elevator.Spring tab deflection is proportional to the control column force and,therefore, to the aerodynamic load imposed on the elevator. At lowspeeds, the spring tab remains in the neutral position. At high speeds,where the aerodynamic load is greater, the tab functions as an aid inmoving the elevator.

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PITCH TRIM SYSTEM

GENERALPitch trim is accomplished by an electrically-actuated movablehorizontal stabilizer. The system may be either automatically ormanually commanded. In both cases, the pitch trim signal is sent to theHorizontal Stabilizer Control Unit (HSCU) channels, which afterprocessing it, command the electric motor in the Horizontal StabilizerActuator (HSA).

SYSTEM COMPONENTS

Horizontal Stabilizer Control Unit (HSCU)The Horizontal Stabilizer Control Unit (HSCU) is located in the rearelectronic compartment at the rear fuselage. It incorporates twoidentical control channels, main and backup. These channel operationsare totally independent from each other. If the pitch trim main channelis inoperative, the horizontal stabilizer can still be commanded throughthe backup channel.The HSCU controls the trimming rate (in degrees/second) based uponthe airplane airspeed. The trimming rate reduces as the airspeedincreases. The HSCU also checks the stabilizer surface position.When the Takeoff Configuration Check Button is pressed, if the surfaceis not within the takeoff green band limits, an aural warning message issounded to the crew.

Horizontal Stabilizer Actuator (HSA)The Horizontal Stabilizer Actuator (HSA) consists of anelectromechanical actuator driven by two DC motors. One of themotors is driven by the main control channel of the Horizontal StabilizerControl Unit (HSCU) and the other motor is driven by the backupchannel of the HSCU. Only one motor will be driven at a time.

SYSTEM OPERATIONPitch trim commands may be done manually through the mainswitches on the control wheels or through backup switch on the controlpedestal aft panel and automatically commanded through the autopilotor speed brake actuation.When using the main control wheel trim switches or the backup trimswitch, it is necessary to command both halves simultaneouslybecause, if just one half is commanded, the control unit will not provideany command to the actuator.In the case of activation of any stick shaker, the pitch trim up commandwill be inhibited.

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PITCH TRIM CHANNELS PRIORITY

Command priorities are: LH switch actuation overcomes the RH switchactuation, which, in turn, overcomes the autopilot. There is no prioritywith respect to the actuation of the main pitch trim switches and thebackup pitch trim switches, the first being commanded takes priority.The main and backup pitch trim switches should not be commandedsimultaneously. For the case of a simultaneous command of bothchannels, the message PTRIM MAIN INOP or PTRIM BACKUP INOPwill be displayed on the EICAS, associated to the second switchcommanded. This message will disappear around 4 seconds after thesecond pitch trim switch is released.

PITCH TRIM SYSTEM PROTECTION

Switch Protection

When only one half of the main control wheel trim switch or backuptrim switch is commanded for more than 7 seconds continuously, thecontrol unit will recognize that one half of the switch is failed stuck atthe commanded position and will disregard any other commandcoming from that faulty switch. A TRIM voice message is provided toalert pilots that just one half of switch is being commanded and themessages PTRIM CPT SW FAIL, PTRIM F/O SW FAIL andPTRIM BKP SW FAIL will be displayed on the EICAS.

Runaway Protection

A quick-disconnect button on each control wheel allows disconnectionfrom the entire pitch trim system. In case of a runaway horizontalstabilizer, the button must be kept pressed until a definitedisengagement is accomplished through the cutout buttons on thecontrol pedestal.

Inadvertent Actuation Protection

A continuous command of any trim switch is limited to 3 seconds, evenif the trim switch is pressed longer than 3 seconds. As a result, whenmanually actuating the trim, it is necessary to release the switch after a3-second actuation, then actuate it again to continue the trimcommand. This feature intends to minimize the effects of aninadvertent trim command of the main and backup trim switches orGround Spoiler/Speed Brake Unit. The autopilot command is notlimited in time and has another logic to preclude inadvertent actuation.

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NOTE: For airplanes equipped with an HSCU -5009 MOD.2 or newerand AWU -5 a TRIM voice message is provided to alert pilotsthat the trim switch is being pressed for more than 3 seconds.

HSA Excessive Load Protection

The crew should keep the airplane trimmed to avoid excessive loadson the Horizontal Stabilizer Actuator (HSA), especially after takeoff.High loads on horizontal stabilizer may stall the HSA, inducing atemporary loss of pitch trim command. If the pitch trim switches areactuated during the period when the Horizontal Stabilizer Actuator isstalled, the message PTRIM MAIN INOP or PTRIM BACKUP INOP willbe displayed on the EICAS. The message will disappear if the trimswitch is released or any horizontal stabilizer motion is detected. If thetrim switches are actuated for a period of time that totalizes 16seconds during the period when the horizontal stabilizer actuator isstalled, the control unit will switch the associated system (main orbackup) off and the message PTRIM MAIN INOP or PTRIM BACKUPINOP will be permanently displayed on the EICAS.


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EICAS MESSAGES

TYPE MESSAGE MEANINGPTRIM MAIN INOP Pitch trim main system is

inoperative, or

Quick Disconnect button iskept pressed for more than11 seconds. This messagewill disappear after thebutton is released, or

Main trim switch(es)actuation associated withthe horizontal stabilizerbeing commanded by thebackup switch, or

Main trim switch beingactuated with the HSAstalled.

WARNING PTRIM BACKUP INOP Pitch trim backup system isinoperative, or

Quick Disconnect button iskept pressed for more than11 seconds. This messagewill disappear after thebutton is released, or

Backup trim switchactuation associated withhorizontal stabilizer beingcommanded by the mainchannel, or

Backup trim switch beingactuated with the HSAstalled.

PTRIM CPT SW FAILPilot´s pitch trim switch isinoperative.

CAUTION PTRIM F/O SW FAILCopilot´s pitch trim switch isinoperative.

PTRIM BKP SW FAILPitch trim backup switch isinoperative.

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CONTROLS AND INDICATORSCONTROL STAND

1 - ELEVATOR DISCONNECTION HANDLE− When pulled, disconnects pilot's from copilot's controls.− To pull the handle, the safety lock button must be pressed.

2 - ELEVATOR DISCONNECTION LIGHT− Illuminates to indicate that the elevator mechanism is

disconnected.

CONTROL STAND

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CONTROL WHEEL

1 - PITCH TRIM SWITCH (spring-loaded to neutral)− Allows trimming the airplane when the autopilot is not engaged.

The trim switch is a 3-position (UP/OFF/DN) rocker switch.− Operating the switch while the autopilot is engaged causes the

autopilot to disengage.− It is divided into two segments, which have to be actuated

together to provide command.

2 - QUICK-DISCONNECT BUTTON (momentary action)− When pressed, disconnects all trim systems.

CONTROL WHEEL


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CONTROL PEDESTAL AFT PANEL

1 - PITCH TRIM MAIN SYSTEM CUTOUT BUTTON (safety guarded)− Cuts out (pressed) or enables (released) the Main Pitch Trim

system.− A striped bar illuminates inside the button to indicate that it is

pressed.− Autopilot is not available.

2 - PITCH TRIM BACKUP SYSTEM CUTOUT BUTTON (safety guarded)− Cuts out (pressed) or enables (released) the Backup Pitch Trim

system.− A striped bar illuminates inside the button to indicate that it is

pressed.− Autopilot is available.

3 - PITCH TRIM BACKUP SWITCH (spring-loaded to neutral)− Pressed forward or backward actuates the pitch trim through the

backup channel.− Operation of the switch while the autopilot is engaged causes

the autopilot to disengage.− It is divided into two segments, which have to be actuated

together to provide command.

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EICAS INDICATION

1 - PITCH TRIM INDICATION

− A green pointer moving on a white vertical scale represents theamount of pitch compensation.

− Trim position is indicated digitally in a white box.

− The letters UP or DN are presented above the box to indicatethat the airplane is trimmed up or down.

− Scale ranges from 4° nose down (bottom of scale) to 10° noseup (top of scale). Every thick mark on the scale represents3.5° of pitch.

− A green band is provided on the analog scale from 4° to 8° noseup to indicate the allowable takeoff position range for thehorizontal stabilizer.

NOTE: Due to the system’s resolution, it’s possible to have the digits,box and pointer turning amber, in spite of the fact that the pitchtrim indication is displayed at 4º or 8º. The trim setting colordisplayed on the EICAS depends on the horizontal stabilizersurface position. For the unit 8 displayed on the EICAS thesurface position can be between 7.7° and 8.7° going upwardand between 8.3° and 7.3° going downward. The color changewould occur when the surface position is 8.1°. For this reason,when setting pitch trim to 8, first select 7. Then, increaseslowly and stop trimming immediately when the value 8 isdisplayed. For the unit 4 displayed on the EICAS, the surfaceposition can be between 3.7° and 4.7° going upward andbetween 4.3° and 3.3° going downward. The color changewould occur when the surface position is 3.9°. For this reason,when setting pitch trim to 4, first select 5. Then, decreaseslowly and stop trimming immediately when the value 4 isdisplayed. This procedure prevents to set the trim at the top orbottom of the green band in order to avoid the possibility ofencountering takeoff config warnings.

− In the event of a pitch trim miscomparison, the pointer, digitalvalue, and the direction indication are removed from display.

− If the pitch trim is out of the green band and the airplane is onthe ground, the pointer and digital indications will turn amber.

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− If the airplane is on the ground and any thrust lever angle isabove 60° and pitch trim is outside the green band, the digits,box, and pointer turn red and sound the aural warningTAKEOFF TRIM and the EICAS message NO TAKEOFFCONFIG is also displayed.

EICAS INDICATIONS


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ROLL CONTROL

Roll control is provided by hydraulically-actuated ailerons controlled byeither control wheel.


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AILERON CONTROL SYSTEMThe ailerons are positioned by the pilot´s control wheels, which arelinked together by a torque tube and cables to supply mechanical inputto two separate hydraulic actuators.

Each aileron actuator is supplied by both hydraulic systems. Eitherhydraulic system is capable of providing full power control. Ifnecessary, each hydraulic system supply can be shut off, by means ofa button installed on the overhead panel. In case of loss of bothhydraulic systems, rotation of the pilot´s control wheels mechanicallypositions the ailerons.

In case of jamming of either aileron, both panels may be disconnectedthrough a handle located on the control pedestal. This procedure willrelease the free aileron from its jammed counterpart allowing the freepanel to be commanded. When disconnected, an amber lightilluminates on the control stand. Controls cannot be reconnectedduring flight, requiring maintenance action.

An autopilot servo is installed on the left side of the torque tube. Theroll trim servo and the artificial feel unit are installed on the right side ofthe torque tube. In case of system disconnection, the artificial feel unitwill actuate on the right aileron only and the autopilot must not be used.The artificial feel unit is provided to give pilots a aerodynamic loadfeedback imposed on the aileron surface.

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ROLL TRIM SYSTEM

Roll trim is performed by relocating the aileron’s neutral position. It isprovided through an electromechanical actuator linked to the artificialfeel unit and commanded through a switch on the control pedestal aftpanel. If the aileron trim switches are activated with the autopilotengaged, the aileron neutral point is repositioned. When the autopilotis disengaged, the ailerons move to the repositioned aileron neutralpoint.

A continuous command of the roll trim switch is limited to 3 seconds,even if the trim switch is pressed longer than 3 seconds. As a result,when manually actuating the trim, it is necessary to release the switchafter a 3-second actuation, then actuate it again to continue the trimcommand. This feature intends to minimize the effects of aninadvertent trim command failure.

When using the roll trim switch, it is necessary to command bothsegments simultaneously since, if just one segment is commanded,the control unit will not provide any command for the actuator.

A quick-disconnect button installed on the control wheels allows, whilekept pressed, to disconnect the roll trim.

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EICAS MESSAGES

TYPE MESSAGE MEANING

CAUTIONAIL SYS 1 (2) INOP Aileron actuation through

hydraulic power is inope-rative.

CONTROLS AND INDICATORS

FLIGHT CONTROLS PANEL

1 - AILERON SHUTOFF BUTTON− Enables (pressed) or disables (released) the associated aileron

hydraulic actuator.− A striped bar illuminates in the button to indicate that it is

released.

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CONTROL STAND

1 - AILERON DISCONNECTION HANDLE

− When pulled, disconnects pilot's from copilot's controls.− To pull the handle, the safety lock button must be pressed.

2 - AILERON DISCONNECTION LIGHT

− When the striped bar is illuminated, indicates that the ailerondisconnection mechanism is actuated.

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1 - ROLL TRIM SWITCH (spring-loaded to neutral)

− Pressed left or right actuates the roll trim to roll left or right.


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EICAS INDICATIONS

1- ROLL TRIM POSITION

− Indicated by a green pointer moving on a white semicircle scale.− Center of the scale is zero trimming.− Each mark represents 50% of trimming range for the associated

side.

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YAW CONTROLYaw control is provided through hydraulically-powered rudders, whichmay also be mechanically commanded. A yaw trim system assists inmoving and holding the rudder in the desired position.


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RUDDER CONTROL SYSTEM

Directional control about the yaw axis is provided by two in-tandemrudders. Forward rudder is driven by the control system, while the aftrudder is linked to the forward rudder and deflected as a function offorward rudder deflection. Either set of rudder pedals will position therudder through a Power Control Unit (PCU). The mechanical control isfully duplicated, consisting of cables running from the pedals in thecockpit to the rear fuselage, where the PCU is commanded to positionthe forward rudder. The rudder can also be commanded through theautopilot.

The rudder PCU is a dual hydraulic unit, simultaneously powered byboth hydraulic systems. Each PCU hydraulic circuit controls thehydraulic power to one respective rudder actuator. Therefore, therudder system is divided into Rudder System 1 and Rudder System 2.The PCU also incorporates an artificial feel device that provides thepedals with an artificial feel of the aerodynamic load imposed on therudder.

Rudder System 1 and/or Rudder System 2 may be either manually orautomatically shut off. The manual shut off operation is providedthrough the Rudder Shutoff Buttons, located on the Overhead Panel.The automatic shut off operation is provided through the speed switchand through the hardover protection function.

When operating under mechanical mode the aerodynamic loads on therudder are directly transmitted to the pedals and, therefore, to thepilots. Since no rudder hydraulics control is available, artificial feel andtrim functions will also not be available. Some characteristics can beobserved:− greater control forces;− sluggish response of rudder to pedals inputs;− backlash of rudder pedals around neutral position when changing the

force application from one to the other pedal.

If either or both rudder systems are inoperative, caution messages arepresented on the EICAS.

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AUTOMATIC SHUTOFF THROUGH THE SPEED SWITCH

During normal operation both systems are powered at speeds below135 KIAS. Above 135 KIAS, Rudder System 1 is automatically shut off.

If the automatic shut off fails to shut off a system above 135 KIAS, acaution message is presented on the EICAS. In this case, it isnecessary to manually shut off system 1 or 2, according to thechecklist.

If Rudder System 2 hydraulic power supply fails, Rudder System 1automatically takes over the rudder and an associated cautionmessage is presented on the EICAS.

RUDDER HARDOVER PROTECTION

The rudder hardover protection function automatically selects themechanical reversion mode as a function of pedal input force, rudderdeflection, and airplane engine operation (two or single-engineoperation). This feature is applicable in the case of a runaway rudderand a caution message is presented on the EICAS.The rudder systems are automatically shut off if all conditions beloware met simultaneously:− Rudder deflected above 5° ± 1°.− Force above 59 kg (130 lb) on the pedal to counteract rudder

deflection.− Both engines running above 56% N2.

CAUTION: DO NOT RESET THE RUDDER SYSTEMS IF THEMECHANICAL REVERSION MODE WAS RESULTANTOF HARDOVER PROTECTION ACTIVATION.

If mechanical reversion mode was not resultant of hardover protection,a reset function is available on the Overhead Panel, by pressing bothRudder Shutoff Buttons off and on again.

The following remarks are applicable to the rudder hardover protection:

• The signal from the Pedal Spring-Loaded Cartridges to shut off therudder systems are applicable only if the pilots are applying force toone side with the rudder deflected above 5° ± 1° to the oppositeside. If pilot command input and the rudder deflection are in thesame direction, the system will not be shut off, regardless of howstrong the pilot input.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


5 Code

01

• The rudder hardover protection is disabled at high airspeeds, whenthe rudder deflection authority decreases to less than 5°.

• During single-engine operation, when the rudder system is moresignificantly required, the shut off function is disabled.

• If a disagreement between FADECs from the same engine occurs,rudder hardover protection is deactivated and a caution message ispresented on the EICAS.

RUDDER DEFLECTION

AIRPLANES UNDER CTA AND FAA CERTIFICATION

The rudder’s main control primary stops, limit rudder deflection at± 15° on ground or in flight.

AIRPLANES UNDER JAA CERTIFICATION

These airplanes are equipped with movable rudder primary stops,which provide two different ranges of rudder deflection:− On ground: maximum rudder deflection is ± 15°.− In flight: maximum rudder deflection is ± 10°.

The Movable Rudder Primary Stops System comprises a hydraulicactuation system, which operates according to the air/ground logic andwill limit rudder deflection to 10° in the extended position and to 15° inthe retracted position.

An amber indication light is provided on the main panel to alert thecrew in case of a disagreement between the actuator position and theair/ground condition.


MANUAL

2-13-20 Page

6 Code


YAW TRIM SYSTEM

Yaw trim is accomplished by an electromechanical actuator, whichreceives signals from the yaw trim knob.

A continuous command of the yaw trim knob is limited to 3 seconds,even if the trim knob is actuated longer than 3 seconds. As a result,when manually actuating the trim, it is necessary to release the knobafter a 3-second actuation, then actuate it again to continue the trimcommand. This feature intends to minimize the effects of aninadvertent trim command failure.

Yaw trim position is presented on EICAS display.A quick-disconnect button installed on the control wheels allows, whilekept pressed, disconnecting the yaw trim.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


7 Code

01

YAW TRIM SCHEMATIC


MANUAL

2-13-20 Page

8 Code


EICAS MESSAGES

TYPE MESSAGE MEANINGRUDDER SYS 1 INOP Rudder System 1 is

inoperative. Message ispresented under thefollowing conditions:− Below 135 KIAS.− Above 135 KIAS if

airspeed of both ADC’s isinvalid.

RUDDER SYS 2 INOP Rudder System 2 is inop-erative.

RUDDER SYS 1–2 INOP Both Rudder Systems areinoperative.

CAUTION RUDDER OVERBOOST Both rudder systemshydraulic actuators arepressurized above 135KIAS.

RUD HDOV PROTFAIL − Disagreement betweenboth FADECs of a sameengine.

− Rudder position micro-switches indicate rudderto right and left simul-taneously.

RUD STOP DISAGREE (*) The rudder’s movable stoppresents disagreement: 15°in flight or 10° on ground.

(*) Applicable to airplanes operating under JAA certification and notequipped with rudder movable stops indication light.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


9 Code

01



1 - RUDDER SHUTOFF BUTTON− Enables (pressed ) or disables (released) the associated rudder

hydraulic actuator.− A striped bar illuminates in the button to indicate that it is

released.



MANUAL

2-13-20 Page

10 Code

01 REVISION 4


1 - YAW TRIM KNOB (spring-loaded to neutral)

− Rotated clockwise or counterclockwise actuates the yaw trim,right or left .


AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


11 Code

01

MAIN PANEL

1 - MOVABLE RUDDER STOPS INDICATION LIGHT(APPLICABBLE TO AIRPLANES OPERATING UNDER JAACERTIFICATION)

− Color: amber

− Illuminates to indicate an incorrect position of at least onehydraulic actuator of the movable rudder stops system, asfollows:- Airplane in flight with movable rudder stops at 15° position.- Airplane on ground with movable rudder stops at 10°position.

− A time delay of 5 seconds is provided to prevent fault indicationduring transient.

NOTE: For some airplanes, the indication light will be replaced by theEICAS message RUD STOP DISAGREE.

MAIN PANEL


MANUAL

2-13-20 Page

12 Code


EICAS INDICATIONS

1- YAW TRIM POSITION

− Indicated by a green pointer moving on a horizontal scale.− Center of the scale is zero trimming.− Each mark represents 50% of trimming range for the associated

side.

EICAS INDICATIONS

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


1 Code

01

GUST LOCK SYSTEMA gust lock system is provided to lock the elevator to avoid damage toelevator components in the case the aircraft is subject to strong gustson the ground. The aileron and rudder surfaces do not need to bemechanically locked since their actuation systems naturally damp anyundesired movement.

ELECTROMECHANICAL GUST LOCK SYSTEM

The electromechanical gust lock can be identified by a yellow andblack striped safety lock device on the control pedestal with theinscription ELEC GUST LOCK, and by two indication lights on theglareshield panel.

The electromechanical gust lock acts directly on the elevator panels,preventing them from moving. Basically, the system is composed oflocking pins driven by an electromechanical actuator, which iscommanded by the gust lock lever. Gust lock system operation(locking and unlocking) is possible on the ground only. Once airborne,the system is deenergized to prevent gust lock lever movement andinadvertent actuation.

The gust lock indication lights located on the glareshield panelilluminate to indicate the unlocking cycle or when a failure in thesystem occurs or when it is pressed for test. For airplanes Post-Mod.SB 145LEG-27-0011 or with an equivalent modification factoryincorporated, when the TLA is higher than 59° and the gust locksystem is still locked, the light will illuminate indicating that anunlocking cycle has initiated.

The system is fed by DC Bus 2 and has a dedicated circuit breaker,located on the overhead circuit breaker panel.


MANUAL

2-13-25 Page

2 Code

01 ORIGINAL

LOCKING OPERATION

To lock the elevator proceed as follows:

A. Pull the control column backwards to any position from neutral tofull nose up.

B. Lift the safety lock device (1) and move the gust lock lever from theunlocked to the locked position (2).

C. Push the control column fully forward until the control columnmovement is restricted. Locking is completed.

In locked position, the thrust levers are prevented from moving beyondthe thrust setting needed for ground maneuvering. However, the gustlock lever was designed to allow extra travel for one of the thrustlevers.

NOTE: During the locking operation, indication lights remain off.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


3 Code

01

TO LOCK:


MANUAL

2-13-25 Page

4 Code


UNLOCKING OPERATION

To unlock the elevator proceed as follows:

A. Lift the safety lock device (1) and move the gust lock lever to itsintermediate detented position (2).

B. At this position, the locking pins are commanded to open and theelevators will be unlocked after approximately 8 seconds. Theindication lights will illuminate during the unlocking cycle, remainingoff after that.

After the indication lights go off, pull the control column backwardsto any position from neutral to full nose up.

C. Lift the safety lock device (3) and pull the gust lock lever from theintermediate position to its full forward inflight resting position (4),completing the unlocking cycle.

NOTE: Gust lock lever command from the intermediate to the unlockedposition is not possible prior to pulling column rearward.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


5 Code

01

TO UNLOCK:


MANUAL

2-13-25 Page

6 Code

01 REVISION 5


GLARESHIELD PANEL

GUST LOCK INDICATION LIGHTS

− Color: amber− Illuminates during the unlocking cycle to indicate that the locking

pins were commanded to unlock the elevator surfaces.− Illuminates in case of failure.− Illuminates when it is pressed.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


7 Code

01

CONTROL STAND

GUST LOCK LEVER

− Actuated backward, locks both elevator and thrust control levers.− The safety lock has to be lifted to move the lever.

CONTROL STAND


MANUAL

2-13-25 Page

8 Code



AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


1 Code

01

FLAP SYSTEMThe flaps are electrically operated, consisting of two double-slotted flappanels installed to each wing.


MANUAL

2-13-30 Page

2 Code


FLAP SYSTEM OPERATION

The Flap Selector Lever provides four detent settings at 0°, 9°, 22° and45° positions. Intermediate positions cannot be selected. When anyposition is selected, the selector lever signals to the Flap ElectronicControl Unit (FECU) to move the flap panels. The FECU also monitorssystem failures and flap position, sending signals to the EICAS andother related systems.

Flap Power and Drive Unit (FPDU) drive the flap panels. The FPDU isa gearbox with two electric motors connected to that unit. Each motoris controlled by the FECU through one independent channel. Bothmotors drive all the flap actuators through flexible shafts. If a motor, orits associated FECU control channel, or associated velocity sensor ortransmission brake fail, the affected channel is disengaged and itsassociated motor actuation is interrupted. The remaining motor candrive all flap panels at half speed. An EICAS message is presented toindicate that flaps are being moved at a lower speed. If both motors orcontrol channels fail, an EICAS message is presented to indicate thatthe system is inoperative.

Flap actuators are torque-limited to safeguard structure againstexcessive loading should flaps or actuators jam. Velocity sensorsinstalled at the end of the flexible shafts detect panels asymmetry. Insuch cases, the system is disabled.

On the ground a protection circuit prevents flap movement when theairplane is energized and a disagreement is detected between flapposition and flap selector lever. To override such protection, it isnecessary to lift up and release the flap selector lever.

Two switches on the Flap Selector Lever send signals to the LandingGear Warning System to alert pilots any time the airplane is in alanding configuration and the gear legs are not locked down.

Flap position is shown on the EICAS display. There are also flap markson the wing trailing edge, indicating 9° and 22°, which becomes visiblewhen flap moves to those positions.

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MANUAL

2-13-30 Page

4 Code


EICAS MESSAGES


CAUTION FLAP FAILBoth flap channels are ino-perative or flap asymmetryis 2° or more.

ADVISORY FLAP LOW SPEEDOne flap channel isinoperative.

FLAP AURAL WARNING (TAKEOFF FLAPS)

If any thrust lever angle is above 60° and the flaps are not in theappropriate takeoff position while the aircraft is on the ground, theTAKEOFF FLAPS aural warning and the EICAS configuration warningwill be activated.



1 - FLAP SELECTOR LEVER

− Moved to the detent positions, selects each discrete flap position.− To move the lever it is necessary to pull it up.− Intermediate positions are not enabled.

NOTE: The flap position 18° can not be selected.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


5 Code

01



MANUAL

2-13-30 Page

6 Code

01 REVISION 4

EICAS INDICATIONS

1- FLAPS POSITION

− Ranges from 0° to 45°, with discrete indication on 0°, 9°, 22° and45°.

− Colors:− Box: white.− Digits: - green (except 0, which is white).

- changes to a green dash when flaps are in transit.− In-transit flap position is replaced by the actual flap position if

flap fails.− If data is invalid, digits are replaced by amber dashes and box

becomes amber.

NOTE: The Flap position can be seen on RMU display.

EICAS INDICATIONS

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


1 Code

01

SPOILER SYSTEM

Spoiler system consists of speed brake and ground spoilersubsystems. Speed brakes allow increased descent rate and assist indecelerating the airplane. Ground spoilers destroy lift, thus providingbetter braking effectiveness.

Spoilers are electrically commanded and hydraulically actuated. ASpoiler Control Unit is responsible for permitting the spoiler panels toopen or not. Four spoiler panels are provided, two per wing surface.The outboard spoilers provide both speed brake and ground spoilerfunctions, while the inboard spoilers provide only a ground spoilerfunction. The actuation of either subsystem is fully independent.


MANUAL

2-13-35 Page

2 Code

01 REVISION 6

GROUND SPOILER

The Spoiler Control Unit (SCU) automatically performs ground spoileropening, without pilots' interference. The SCU enables the groundspoilers to open whenever the following conditions are met:

− Airplane on the ground.− Main landing gear wheels running above 25 kt.− Both engines thrust lever angles set to below 30° or both engines N2

below 56%.

If any of those conditions is not met, the ground spoilers will not open.A status indication is presented on the EICAS to indicate that thespoilers are open or closed. If a failure is detected, a caution messageis presented on the EICAS.

SPEED BRAKE

When speed brake is commanded with autopilot engaged, the autopitch trim is provided through the autopilot; when the autopilot is notengaged the Spoiler Control Unit provides the auto pitch trimcommand.

The speed brakes will open when the speed brake lever is set to openand the following conditions are met:

− Thrust lever angle of both engines set to below 50°.− Flaps at 0° or 9°.

If the speed brake lever is commanded to the OPEN position and anyof the speed brake open condition is not met, the speed brake panelsare kept closed and a caution message is presented on the EICAS. Ifthe speed brake panels are open and any of the speed brake opencondition is not met, the speed brake panels automatically close andan EICAS message is presented. In both cases, the speed brake levermust be moved to the CLOSE position to remove the EICAS message.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


3 Code

01

SPOILER SYSTEM SCHEMATIC


MANUAL

2-13-35 Page

4 Code


EICAS MESSAGES


CAUTION

SPOILER FAIL Any spoiler panel openinadvertently, failed to openor any failure in the inputsignals.

SPBK LVR DISAGREE Speed Brake Levercommanded to OPEN butopening logic is notsatisfied.

SPOILER AURAL WARNING (TAKEOFF SPOILERS)

If any thrust lever angle is above 60° and any spoiler/speed brakepanel is deployed, the spoiler OPN label turns red and the TAKEOFFSPOILERS aural warning and the EICAS configuration warning will beactivated.


CONTROL STAND

1 - SPEED BRAKE LEVER

− Actuated to the OPEN position commands outboard spoilerpanels to open, provided enabling conditions are met.

AIRPLANEOPERATIONS

MANUAL

FLIGHT CONTROLS


5 Code

01

CONTROL STAND


MANUAL

2-13-35 Page

6 Code


EICAS INDICATIONS

1- SPOILERS INDICATION

− Displays OPN when any of the surfaces are open, or CLD whenall of the surfaces are closed.

− Colors:− Box: white.− CLD: white.− OPN: - green in normal condition.

- red if any surfaces are open during takeoff.

EICAS INDICATIONS

AIRPLANEOPERATIONS

MANUAL

OXYGEN


1 Code

01

SECTION 2-16

OXYGENTABLE OF CONTENTS

Block Page

General .............................................................................. 2-16-05 ..01Flight Crew Oxygen............................................................ 2-16-10 ..01

EICAS Message............................................................ 2-16-10 ..05ECS Page on MFD ....................................................... 2-16-10 ..06Crew Mask Stowage Boxes .......................................... 2-16-10 ..08Crew Mask .................................................................... 2-16-10 ..09Controls and Indicators (EROS Mask).......................... 2-16-10 ..10Smoke Goggles ............................................................ 2-16-10 ..12

Passenger Oxygen............................................................. 2-16-15 ..01EICAS Message............................................................ 2-16-15 ..04ECS Page on MFD ....................................................... 2-16-15 ..04Passenger Dispensing Unit and Mask .......................... 2-16-15 ..06Controls and Indicators ................................................. 2-16-15 ..08

Portable Oxygen ............................................................... 2-16-20 ..01 Portable Oxygen Cylinder ............................................. 2-16-20 ..01 Protective Breathing Equipment.................................... 2-16-20 ..04Minimum Oxygen Pressure for Dispatch ........................... 2-16-25 ..01

Flight Crew Oxygen Subsystem.................................... 2-16-25 ..01Passenger Oxygen Subsystem..................................... 2-16-25 ..01Portable Oxygen Cylinder ............................................. 2-16-25 ..01Oxygen Pressure Correction Chart............................... 2-16-25 ..02Oxygen Consumption Chart.......................................... 2-16-25 ..04

OXYGEN AIRPLANEOPERATIONS

MANUAL

2-16-00 Page

2 Code

01 ORIGINAL


AIRPLANEOPERATIONS

MANUAL

OXYGEN


1 Code

01

GENERAL

The oxygen system supplies, in case of cabin decompression,protective and supplemental oxygen for the crew and onlysupplemental oxygen for the passengers. It is a conventional, high-pressure gaseous type system, in which the oxygen is stored in acylinder at high pressure and distributed at low pressure to the masks.

The system is composed of three subsystems that operateindependently: The Flight crew oxygen subsystem, the Passengeroxygen subsystem and the portable oxygen subsystem.

The first two subsystems are monitored so that all the necessaryparameters are informed to the flight crew and flight attendants.


MANUAL

2-16-05 Page

2 Code



AIRPLANEOPERATIONS

MANUAL

OXYGEN


1 Code

01

FLIGHT CREW OXYGENThe flight crew oxygen subsystem employs a 50-cu.ft cylinder, installedon the right side of the cockpit/passenger cabin partition, in which theoxygen is stored at high pressure (1850 psi) to supply the cockpit crewmembers (Pilot, Copilot and Observer) masks.

The system is protected from overpressurization by a safety disclocated on the lower right side of the aircraft’s nose. Discharge throughthe safety disc may be visually verified when the discharge indicator(green disc) has been blown out.

If the cylinder pressure drops below 400 psi (for airplanes equippedwith EICAS prior to version 20.5) or 450 psi (for airplanes equippedwith EICAS version 20.6 and on), a caution message is presented onthe EICAS.

The cylinder is provided with an integrated shutoff/regulator valve, thatcontrols oxygen outlet pressure. The regulator valve at the ON positionsupplies the crew distribution lines at a low-pressure rate (70 psi). Arelief valve opens if the pressure exceeds 90 psi.

On the left side of the oxygen service panel, located on the right side ofthe front fuselage, there is a crew charging valve and a pressuregauge to check oxygen quantity. Some airplanes may have a factoryinstalled removable panel located behind the copilot’s seat thatprovides access to the oxygen cylinder and its replacement. Thecylinder pressure is also indicated on the MFD (ECS page).

The cockpit is provided with a quick donning diluter/demand-type mask(or a pressure/demmand type for 41000 ft operation), available insidemask stowage boxes adjacent to each crew station, and a smokeprotection kit, which consists of two smoke goggles to be used by thepilot and copilot with the diluter/demand masks (or pressure/demmandmasks for 41000 ft operation).

There are two different types of flight crew oxygen masks, as follows:− Flight crew oxygen mask with dilution demand regulator type,

required for airplanes with altitude ceiling of 39000 ft;− Flight crew oxygen mask with pressure demand regulator type,

required for airplanes with altitude ceiling of 41000 ft.

NOTE: The two types of flight crew oxygen masks presents nodifferences in its size and operation.


MANUAL

2-16-10 Page

2 Code

01 REVISION 7

FLIGHT CREW OXYGEN SYSTEM SCHEMATIC

AIRPLANEOPERATIONS

MANUAL

OXYGEN


3 Code

01

CREW OXYGEN CYLINDER


MANUAL

2-16-10 Page

4 Code


OXYGEN SERVICE PANEL

AIRPLANEOPERATIONS

MANUAL

OXYGEN


5 Code

01

EICAS MESSAGE


CAUTION

CREW OXY LO PRESS Crew oxygen cylinderpressure below 400 psi (forairplanes equipped withEICAS prior to version 20.5)or 450 psi (for airplanesequipped with EICASversion 20.6 and on).Remaining oxygen sufficientfor approximately 12minutes for pilot, copilot,and observer.


MANUAL

2-16-10 Page

6 Code

01 REVISION 7

ECS PAGE ON MFD

1 - ANALOGIC OXYGEN PRESSURE INDICATION

AIRPLANES EQUIPPED WITH EICAS PRIOR TO VERSION 20.5

Pointer:− Green between 410 to 1850 psi.− Amber between 250 to 400 psi.− Red between 0 to 240 psi.

AIRPLANES EQUIPPED WITH EICAS VERSION 20.6 AND ON


2 - DIGITAL OXYGEN PRESSURE INDICATION


− Digits are green between 410 to 1850 psi.− Digits are amber between 250 to 400 psi.− Digits are red between 0 to 240 psi.

(Ranges from 0 to 1850 psi, with a resolution of 10 psi).




AIRPLANEOPERATIONS

MANUAL

OXYGEN


7 Code

01

CREW ECS PAGE ON MFD


MANUAL

2-16-10 Page

8 Code

01 REVISION 7

CREW MASK STOWAGE BOXES

The crew mask stowage boxes are directly connected to the oxygendistribution line and to the communications system. The pilot andcopilot boxes incorporate a shutoff valve, which keeps the maskregulator unpressurized while in the stowed position.

When the box doors are opened, the shutoff valve is brought to theopen position, thus allowing oxygen to flow to the mask.

After the mask has been taken out of the stowage box, the doors canbe closed without interrupting oxygen supply to the mask. To stop theoxygen flow, it is necessary to close the left door and activate theTest/Shutoff Sliding Control.

Pilot and copilot mask stowage boxes are also provided with a flowindicator.

NOTE: The observer’s mask stowage box is not provided withTest/ Shutoff Sliding Control (EROS Mask) and, although themask is permanently pressurized, oxygen will flow only ondemand.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


9 Code

01

CREW MASK

The crew mask is a quick donning oro-nasal mask assembly that cansupply breathing protection up to 39000 ft or 41000 ft and allowsoxygen flow on demand or under pressure, as required.

The mask is provided with an automatic oxygen dilution system thatsupplies pure oxygen at a cabin altitude of over 33000 ft. It can also bemanually selected to the 100% position to provide pure oxygen at allaltitudes or to the EMERGENCY position to maintain in the ventingorifice 100% pure oxygen (or positive pressure if it is an oxygen maskwith pressure demand regulator type).

The quick donning operation is as follows:− Hold the mask with one hand by the mask regulator and the

inflation control valve (red ears).− Pull the mask out of the box.− Press the inflation control valve (red ears) firmly. The harness

inflates rapidly, and takes a round shape large and rigid enoughto allow the user to don it quickly.

− Release the regulator ears. The harness will then deflate,securing the mask to the user's face.

NOTE: The EROS Mask is provided with two red ears. One inflatesthe harness when pressed and the other is fixed.

The oxygen mask with dilute demand regulator type (applicable toairplanes with altitude ceiling of 39000 ft) is equipped with a flowindicator only for the observer mask. The flow indicator for the pilot andcopilot is located in the respective crew mask stowage boxes.

The oxygen mask with pressure demand regulator type (applicable toairplanes with altitude ceiling of 41000 ft) is equipped with a flowindicator for the pilot, copilot and observer.


MANUAL

2-16-10 Page

10 Code

01 REVISION 7

CONTROLS AND INDICATORS (EROS MASK)

MASK STOWAGE BOX/CREW MASK

1 - TEST/SHUTOFF SLIDING CONTROL (spring-loaded in the pilotand copilot boxes only)− When pressed, with the mask stowed, allows testing of the

oxygen mask. Flow indicator turns yellow for a short time. TheOXY ON flag appears on the lid face.

− When pressed, with the mask not stowed and the left doorclosed, shuts off oxygen to the mask. The OXY ON flagdisappears on the lid face.

2 - OXY ON FLAG (white)− Appears when the box shutoff valve is open and oxygen is

supplied to the mask.

3 - FLOW INDICATOR (pilot and copilot boxes only)− A yellow star appears when oxygen is flowing.

4 - VENTING VALVE CONTROL (pilot and copilot masks only)− When actuated forward, opens the venting valve.− A red band is visible to indicate that the control is actuated.

5 - HARNESS INFLATION CONTROL VALVE (red ear)− When pressed, inflates the harness and allows mask donning.

6 - FLOW INDICATOR− The black shutter disappears when pressure is applied to the

mask.

7 - TEST/EMERGENCY SELECTOR KNOB− When rotated clockwise, 100% oxygen is supplied under

positive pressure at all cabin altitudes. This mode must beselected when using smoke goggles.

− When pressed, tests if the regulator demand mechanismoperates satisfactorily.

8 - NORMAL/100% SELECTORN - Oxygen/air mixture is supplied on demand. Mixture ratio

depends on the cabin altitude. Above 33000 ft, pure oxygenis supplied.

100% - Pure oxygen is supplied at all cabin altitudes on demand (oron positive pressure if it is an oxygen mask with pressuredemand regulator type). This mode must be selected inconjunction with the EMERGENCY position, whenprotective breathing is required.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


11 Code

01

MASK STOWAGE BOX/CREW MASK (EROS MASK)


MANUAL

2-16-10 Page

12 Code

01 REVISION 7(Pages 13 and 14 deleted)

SMOKE GOGGLES

The smoke goggles were designed for use with the crew maskassembly, matching the mask face cone. The venting valve, located onthe mask shell and manually actuated by the user, allows directcommunication between venting orifice and goggles.When mask regulator is selected to emergency position, a meteredoxygen flow will be directed to the goggles’ cavity so as to allowcontinuous venting and preventing any infiltration of harmful gases.

SMOKE GOGGLES

AIRPLANEOPERATIONS

MANUAL

OXYGEN


1 Code

01

PASSENGER OXYGENThe passenger oxygen subsystem employs one or two oxygencylinder(s), installed in the closeout panel near the galley, in which theoxygen is stored at high pressure (1850 psi) to supply the passengeroxygen masks.

The subsystem shares with the crew oxygen subsystem theoverpressure discharge indicator (safety disk), located on the lower leftside of aircraft’s nose. For that reason, in case of over pressurizationthe same indicator blows out and can be visually inspected. If thecylinder pressure drops below 750 psi, a caution message is presentedon the EICAS.

The Passenger Oxygen Control Panel is located on the right lateralconsole, above the copilot’s mask stowage box.

Under normal operating conditions, when the “AUTO” mode is selectedon the passenger oxygen control panel and the cabin altitude reaches14000 ft, the altimetric switch energizes a time delay relay, whichenergizes the passenger oxygen on-off solenoid valve to initiate theoxygen flow and pressurize the oxygen distribution manifold. Thepressure, in the distribution line, activates the pneumatic latch, openingthe door and dropping the masks from their dispensing units. At thesame time, the “OXYGEN” indicator light, on the passenger oxygencontrol panel, as well as the “NO SMOKING” and “FASTEN SEATBELTS” signs in the passenger cabin are turned on.The “MANUAL” selection on the passenger oxygen control panelactivates the system when the automatic system fails or at any time asrequired.

On the right side of the oxygen service panel, are the passengercharging valve and the passenger pressure gauge, which allowsaccess to charge the passenger oxygen cylinder and monitoring ofpassenger oxygen quantity.

NOTE: The addition of a second oxygen cylinder, will not effect orchange any of the controls, indications or safety features of thePassenger Oxygen subsystem.


MANUAL

2-16-15 Page

2 Code

01 ORIGINAL


AIRPLANEOPERATIONS

MANUAL

OXYGEN


3 Code

01

PASSENGER OXYGEN SYSTEM SCHEMATIC


MANUAL

2-16-15 Page

4 Code

01 REVISION 7

EICAS MESSAGE


CAUTIONPAX OXY LO PRESS Passenger Oxygen cylinder(s)

pressure below 750 psi.

ECS PAGE ON MFD

1 - ANALOGIC OXYGEN PRESSURE INDICATION





2 - DIGITAL OXYGEN PRESSURE INDICATION







AIRPLANEOPERATIONS

MANUAL

OXYGEN


5 Code

01

ECS PAGE ON MFD


MANUAL

2-16-15 Page

6 Code


PASSENGER DISPENSING UNIT AND MASK

The passenger oxygen masks dispensing units are located on the rightand left overhead valance panels in the passenger seating area, at theflight attendant station, at the galley and in the lavatory. Each unit isequipped with two continuous flow masks and a lanyard valve permask.

The passenger oxygen mask assembly possesses a reservoirbag, a flow indicator, an oxygen supply tubing and a head strapfor securing the mask to the passenger’s face. In addition, onelanyard and one “PULL” streamer is fitted to each oxygen mask.

In the event of a decompression, the mask and the “PULL” streamerremain attached to the oxygen flow valve by the lanyard Thepassenger must pull the mask to his face or pull the “PULL” streamerto release the lanyard valve pin to obtain oxygen flow.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


7 Code

01

DISPENSING UNITS/PASSENGER MASKS


MANUAL

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01 REVISION 3


PASSENGER OXYGEN CONTROL PANEL

1 - OXYGEN ON INDICATOR LIGHT (white)− Indicates that the passenger oxygen system is ON, and the

distribution manifold pressure is at or above 45 psig.

2 - OXYGEN VLV CLOSED INDICATOR LIGHT (white)− Indicates that the passenger oxygen pressure regulator is “OFF”

or if the pressure at the outlet of the oxygen regulator dropsbelow 50 to 55 psi.

3 - PASSENGER OXYGEN SELECTOR KNOBCLOSED - Disables the automatic deployment of passenger

masks. Also resets oxygen ON indicator and passengercabin signs after system activation either on automaticor manual mode.

AUTO - Automatically deploys the passenger masks providedthat cabin pressure altitude is above 14000 ft.

MANUAL (momentary position) - Actuates the passenger oxygensystem at any altitude, overriding the altimetric switch,and may be used in case of AUTO mode failure.

AIRPLANEOPERATIONS

MANUAL

OXYGEN

JANUARY 17, 2003 2-16-15 Page

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PASSENGER OXYGEN CONTROL PANEL


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AIRPLANEOPERATIONS

MANUAL

OXYGEN


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PORTABLE OXYGEN

The portable oxygen subsystem includes one portable oxygen cylinderand two Protective Breathing Equipment (PBE).

PORTABLE OXYGEN CYLINDER

The portable oxygen cylinder is located in the passenger cabinentertainment rack and is used for passenger fist-aid therapeutictreatment only. The cylinder, with 11-cu.ft and three flow outlets,provides a constant flow breath for up to 3 passengers or crewmembers. Also, It is equipped with a pressure gauge indicating supply,a high-pressure safety relief device, a carrying strap and an on-offpressure regulator.

To get specified flow, the mask must be connected to one of thecylinder outlets and the handle turned to the “FULL ON” position.

NOTE: This equipment is intended to be used only for aviationapplications and is to be used only by, or under the supervisionof, a pilot or crew member trained and qualified in its use.


MANUAL

2-16-20 Page

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01 ORIGINAL

PORTABLE OXYGEN CYLINDER

AIRPLANEOPERATIONS

MANUAL

OXYGEN


3 Code

01

PORTABLE OXYGEN CYLINDER LOCATION


MANUAL

2-16-20 Page

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01 REVISION 7

PROTECTIVE BREATHING EQUIPMENT (PBE)

The PBE unit is used for respiratory and visual protection againsteffects of smoke and harmful gases while figthing fire in the aircraft. Itis designed to provide breathing protection for a minimun duration of15 minutes.

There are two EROS PBE installed in the airplane, one is installed inthe entertainment equipment rack and the otherone is installed in theLH cockpit/passenger cabin partition, behind the pilot seat.

The device features a hood which encloses the entire head and naintegrated life support unit which uses a solid state oxygen supply(chemical generator).

OPERATION

1 - Take the box, push the spring lock.2 - Pull the box cover upward.3 - Extract the hood and deploy the hood by a brisk downward

movement.4 - Put the hood on head (It can accommodate bears, long hair and

glasses when properly worn).5 - Pull to actuate ring from the unit.

The over-hood gives protection to the head from flames orincandescent objects which may fall from burning structures orinteriors.

The end of autonomy is felt when oxygen flow noise can not be heardanymore. Hood has then to be removed.

CAUTION: THE OXYGEN PRODUCED BY PBE UNIT WILLVIGOROUSLY ACCELERATE COMBUSTION. DO NOTINTENTIONALLY EXPOSE THE PBE UNIT TO DIRECTFLAME CONTACT OR REMOVE IT IN THE IMMEDIATEPRESENCE OF FIRE OR FLAME. DUE TO OXYGENSATURATION OF THE HAIR. DO NOT SMOKE ORBECOME EXPOSED TO FIRE OR FLAME IMMEDIATELYAFTER REMOVING PBE UNIT.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


5 Code

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PROTECTIVE BREATHING EQUIPAMENT LOCATION


MANUAL

2-16-20 Page

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01 REVISION 7

HOOD SCHEMATIC AND STOWAGE - EROS

AIRPLANEOPERATIONS

MANUAL

OXYGEN

JANUARY 17, 2003 2-16-25 Page

1 Code

01

MINIMUM OXYGEN PRESSURE FOR DISPATCH

FLIGHT CREW OXYGEN SUBSYSTEMCrew comprising Pilot and Copilot: 1100 psi.Crew comprising Pilot, Copilot and Observer: 1500 psi.

PASSENGER OXYGEN SUBSYSTEMAirplane equipped with one cylinder: 1730 psi.Airplane equipped with two cylinders (Optional): 1250 psi.

NOTE: The minimum oxygen pressure for dispatch was calculated atan ambient temperature of 21°C (70°F). For othertemperatures, refer to Oxygen Pressure Correction Chart as afunction of the cylinder compartment temperature.

PORTABLE OXYGEN CYLINDERThe minimum portable oxygen cylinder pressure for dispatch is1200 psi (calculated for a maximum utilization period of 30 minutes).


MANUAL

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2 Code


OXYGEN PRESSURE CORRECTION CHART

An Oxygen Pressure Correction Chart is provided for the maintenancepersonnel's use when recharging the oxygen cylinder. Additionally, itmay be used by the crew to check if the oxygen cylinder’s pressure isabove the minimum oxygen pressure for dispatch.

To use the chart for recharging purposes:Enter the chart with the cylinder compartment temperature (cockpittemperature) and go vertically up to the desired pressure at 21°C.From the intersection point, trace to the left to read the indicatedgauge pressure to be attained.

To use the chart for dispatching purposes:Enter the chart simultaneously with the cylinder compartmenttemperature (cockpit temperature) and indicated gauge oxygenpressure (on MFD or oxygen service panel). The intersectiondetermines the oxygen cylinder’s equivalent pressure at 21°C, byinterpolating the two adjacent standard curves.

EXAMPLE

Associated condition:− Crew............................................................PILOT, COPILOT

AND OBSERVER− Indicated gauge pressure............................1600 PSI− Cylinder compartment temperature ............30°C

As the intersection is above the dashed line for the associatedcondition, the airplane may be dispatched.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


3 Code

01

OXYGEN PRESSURE CORRECTION CHART


MANUAL

2-16-25 Page

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01 REVISION 4

OXYGEN CONSUMPTION CHART

The Oxygen Consumption Chart is provided to allow the Flight Crew toknow the remaining number of pre-flight oxygen mask tests availablebefore the oxygen cylinder recharging is necessary. This chart shouldbe used by the maintenance personnel to choose the best moment torecharge the oxygen cylinder.

The Oxygen Consumption chart has been plotted for 21°C (70°F)conditions. For different temperatures, the Oxygen PressureCorrection chart must be used to obtain the pressure at 21°C and thensee what is the number of the remaining oxygen mask tests.

EXAMPLE

Associated condition:− Crew.................................................................PILOT, COPILOT,

AND OBSERVER− Indicated Gauge Pressure ...............................1750 psi− Cylinder Compartment Temperature................30°C

According to the Oxygen Pressure Correction chart, for the associatedconditions, the pressure for 21°C is 1700 psi.

According to the Oxygen Consumption chart, for 1700 psi there areapproximately 22 remaining pre-flight tests before recharging theoxygen cylinder becomes necessary. The airplane’s dispatch beingtherefore allowed.

NOTE: The oxygen consumption chart is used only for the crewoxygen subsystem, since there is no test for the passengeroxygen subsystem.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


5 Code

01

OXYGEN CONSUMPTION CHART


MANUAL

2-16-25 Page

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01 REVISION 7

OXYGEN DURATION TABLES

The oxygen duration tables allow the Fight Crew to estimate the timeavailable for flying above 10000 ft to perform a mountain clearence, forexample, using the emergency oxygen.

The values presented on the following tables were obtainedconsidering the airplane flying at 41000 ft when the decompressionoccurs and an emergency descent, to a terrain clearence altitude, wasapplied (all engines in idle, landing gear down, flaps up and speedbrakes open).

After the time allowed for the terrain clearence there will be sufficientoxygen to procedure the airplane descent to 10000 ft.

AIRPLANES EQUIPPED WITH ONE 77 FT3 OXYGEN BOTTLE

Altitude Crew Members Passengers on board(ft) 0 pax 1 pax 3 pax 5 pax 7 pax 9 pax 11 pax 13 pax 14 pax

15000 with observer 115 115 115 85 66 53 44 38 35

without observer 150 150 117 85 66 53 44 38 35

18000 with observer 107 107 107 85 66 53 44 38 35


NOTE: - The values above are in minutes.- The calculation of the terrain clearence time was based on a

dispatch pressure of 1730 psig or above for crew andpassenger systems. Also it was considered the presence ofone flight attendant in all calculations.

AIRPLANEOPERATIONS

MANUAL

OXYGEN


7 Code

01

AIRPLANES EQUIPPED WITH TWO 77 FT3 OXYGEN BOTTLE


15000 with observer 115 115 115 115 115 115 106 91 86


18000 with observer 107 107 107 107 107 107 106 91 86


20000 with observer 92 92 92 92 92 84 71 61 57




AIRPLANES EQUIPPED WITH ONE 115 FT3 OXYGEN BOTTLE


15000 with observer 115 115 115 110 90 76 65 56 53


18000 with observer 107 107 107 107 90 75 65 56 53


20000 with observer 92 92 92 73 60 50 43 38 35





MANUAL

2-16-25 Page

8 Code

01 REVISION 5

AIRPLANES EQUIPPED WITH ONE 115 FT3 AND ONE 77 FT3

OXYGEN BOTTLES


15000 with observer 115 115 115 115 115 115 115 105 99


18000 with observer 107 107 107 107 107 107 107 105 99


20000 with observer 92 92 92 92 92 92 80 70 66




AIRPLANEOPERATIONS

MANUAL

ICE AND RAINPROTECTION


1 Code

01

SECTION 2-15

ICE AND RAIN PROTECTION

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-15-05 ..01Bleed Air Thermal Anti-Icing System ................................. 2-15-10 ..01

Wing, Stabilizer and Engine Anti-icing Valves Operational Logic............................ 2-15-10 ..03EICAS Messages ........................................................... 2-15-10 ..07

Windshield Heating System ............................................... 2-15-10 ..08EICAS Messages ........................................................... 2-15-10 ..08

Sensor Heating System ..................................................... 2-15-10 ..09EICAS Messages ........................................................... 2-15-10 ..10

Lavatory Water Drain andNipple Heating System............................................... 2-15-10 ..10

Ice Protection Controls and Indicators ............................... 2-15-10 ..11Ice Protection Control Panel........................................... 2-15-10 ..11

Ice Detection System ......................................................... 2-15-15 ..01EICAS Messages ........................................................... 2-15-15 ..02Control and Indicators .................................................... 2-15-15 ..03

Main Panel.................................................................. 2-15-10 ..03Windshield Wiper System (if installed) .............................. 2-15-15 ..04

Windshield Wiper Control Panel .................................... 2-15-15 ..04


AIRPLANEOPERATIONS

MANUAL

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GENERAL

Airplane ice protection system is provided by heating critical ice buildup areas through the use of either hot air or electrical power. Thesystem is fully automatic and under icing conditions, activates theentire protection system (the only exception is the windshield heatingsystem).

The hot air-heated areas are:− Wing and horizontal stabilizer leading edges.− Engine air inlet lips.

The electrically heated areas are:− Windshields.− Pitot tubes, Pitot-static tube, AOA sensors, TAT probes, ADCs and

pressurization static ports.− Lavatory water drain and potable water service nipples.

Two fully independent wiper systems remove rain from thewindshields.

All ice protection systems provide signals to the EICAS formalfunctioning system display.

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BLEED AIR THERMAL ANTI-ICING SYSTEM

The bleed air thermal anti-icing system is supplied with hot air tappedfrom the engines. In the automatic mode, the system is turned onthrough activation of either ice detector. Manually, setting theOVERRIDE Knob to the ALL position activates the system.

Adequate ice protection for the wing and horizontal stabilizer leadingedges and engine air inlet lips is ensured by heating these surfaces.Hot air supplied by the Pneumatic System is ducted through perforatedtubes, known as Piccolo tubes. Each Piccolo tube is routed along thesurface, so that hot air jets flowing through the perforations heats thesurface. Dedicated slots are provided for hot air exhaustion after thehas been surface heated.

During night flights, inspection lights, installed on the wing-to-fuselagefairing, illuminate the wing leading edges, allowing the crew to checkfor ice accumulation.

Each subsystem comprises an anti-icing valve (pressureregulating/shutoff valve). A restrictor limits the airflow rate supplied bythe Pneumatic System. It is monitored by pressure sensors, thatindicate abnormal low and high air pressure conditions. The pressuresensors protect the respective subsystem against either insufficient orexcessive airflow rate.

The wing and stabilizer low pressure protection mode has a redundantdetection by means of a second low pressure sensor on the stabilizersystem and a differential pressure switch (± 2 psi) that compares rootpressure on the left and right half-wing Piccolo tubes.

Air leakage is detected by thermostats installed close to each ductconnection. Low pressure switches provide an additional protectionagainst unacceptable leakage level.

The Piccolo tubes integrity is monitored as follows:

− Horizontal stabilizer: By one differential pressure switch comparingthe left and right Piccolo tubes pressure.

− Half-wing: It depends on the airplane model. By one differentialpressure switch in each Piccolo tube comparing the root and tippressures or, by manometric switches measuring the tip pressureonly.


AIRPLANEOPERATIONS

MANUAL

2-15-10 Page

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01 REVISION 5

Engine ice protection is provided by heating the engine air inlet lip,through the use of non-temperature-controlled hot air tapped directlyupstream of each high stage valve. As the engine air inlet has enoughairflow surrounding the lip when the engine is running, the engine airinlet lip anti-icing system can be operated on the ground normally andwith no limitations. Each engine has its own protection systemindependent of the airplane’s pneumatic system.

The left hand Pneumatic System supplies the horizontal stabilizer anti-icing subsystem. Each half-wing anti-icing subsystem is supplied by itsrespective side of the Pneumatic System.

The bleed air thermal anti-icing system may be deactivated by buttons,located on the overhead panel.

On the ground, the FADEC incorporates an automatic logic to reducethe maximum available thrust to avoid a sudden engine thrust lossduring lift-off, even with the thrust lever set at MAX position.

In flight, the FADEC allows the engines to deliver the maximum ratedthrust to compensate for the effect of the high bleed air consumptionby the wing and horizontal stabilizer thermal anti-icing subsystems.Moreover, the FADEC provides an automatic logic to ensure aminimum available thrust during icing conditions, even during lowthrust setting conditions. This logic is automatically inhibited when thelanding gear is extended, in order to improve the airplane’s rate ofdescent and glide slope path adjusting capability.

The APU bleed air is not hot enough to perform anti-icing functions.Therefore it must not be used for such applications.

A caution message is presented on the EICAS if the thermal anti-icingsystem is turned on during non-icing conditions.

AIRPLANEOPERATIONS

MANUAL



3 Code

01

WING, STABILIZER AND ENGINE ANTI-ICING VALVESOPERATIONAL LOGIC

Since the Bleed Thermal Anti-icing System is supplied by thePneumatic System, it is integrated to the functional logic that providesautomatic control and protection for the system.

The Wing and Stabilizer Anti-icing Valves receive an electrical inputthat open when the following conditions occur:

− The Ice Detection Test Knob is set to 1 or 2, or− The airplane is in-flight or attained a ground speed above 25 knots,

and− The Ice Detection Override Knob is set to ALL, or− The Ice Detection Override Knob is set to AUTO or ENG and any

ice detector is activated.

NOTE: - The Wing and Stabilizer Anti-icing Valves are inhibited fromopening on the ground and at a ground speed below 25knots to prevent structural damage caused by surfaceheating, except during ice detection testing. The icedetection test should not be activated for more than 15seconds.

The Engine Anti-icing Valves receive an electrical input to open whenthe following conditions occur:

− The Ice Detection Override Knob is set to ALL or ENG, or− The Ice Detection Override Knob is set to AUTO position and any

ice detector is activated, or− The Ice Detection Test Knob is set to 1 or 2.

The engine anti-ice system logic has a narrow range between normaloperating pressures and a low pressure value that, if reached, wouldtrigger an E1(2) A/ICE FAIL message on the EICAS. This messagemay be presented in flight whenever the engines are set at low thrustsettings. This message may be cleared increasing the engine anti-icesystem pressure by advancing the thrust levers with Ice DetectionOverride Knob in AUTO. If the message does clear and the relatedEngine Air Inlet OPEN inscription remains illuminated, the system isoperating normally and the flight may be continued.

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EICAS MESSAGESTYPE MESSAGE MEANING

WARNING

ICE COND-A/I INOP Any Bleed Air Thermal anti-icing subsystem notfunctioning properly undericing conditions.

A/ICE LOWCAPACITY

Low pressure conditiondownstream of any wing orstabilizer anti-ice valve orwing pressure asymmetry.

NO ICE-A/ICE ON Any anti-icing valve opened inflight out of icing conditions.

A/ICE SWITCH OFF Any Bleed Air Thermal anti-icing button turned off.

E1 (2) A/ICE FAIL(if applicable)

− Low pressure condition.− Valve failure.− Any switch failure.− Overpressure condition.− Any system failure.

CAUTION

ENG1 (2) A/ICEFAIL(if applicable)

− Low pressure condition (onground or inflight), or

− Disagreement betweenvalve position and systemcommand.

WG1 (2) A/ICE FAIL(if applicable)

− Low pressure condition.− Valve failure.− Any switch failure.− Duct leakage.− Any system activation failure.

WG A/ICE FAIL(if applicable)

− Low pressure condition, or− Disagreement between

valve position and systemcommand, or

− Piccolo tube failure.WG A/ICE ASYMETRY Asymmetrical degradation of

half-wings anti-ice systemsthermal performance.

CAUTION

STAB A/ICE FAIL − Low pressure condition.− Valve failure.− Any switch failure.− Duct leakage.− Any system activation

device failure.

ADVISORY ENG A/ICE OVERPRES Inflight overpressure conditiondetected.


AIRPLANEOPERATIONS

MANUAL

2-15-10 Page

8 Code

01 REVISION 5

WINDSHIELD HEATING SYSTEMThe windshields are electrically heated to prevent ice and fog formationor for deicing and defogging purposes. Due to a higher thermal inertiato bring heat to windshield inner layer, when Descent phase is initiatedthe system must be turned ON to prevent fogging. During all the othersflight phases, the system must be kept OFF except when icingconditions are anticipated or if situation requires.

The outer glass layer has no structural significance but provides arigid, hard and protected surface.

Windshield heating is accomplished through an electric conductive gridembedded in its interlayer, which functions as an electric resistor.

Individual buttons located on the overhead panel control left and rightwindshield heating. Separate power supplies are provided for eachwindshield heating element and its control circuit.

Each windshield element is provided with three temperature sensors.One sensor is used for temperature control and a second sensor isused for overheat protection. A third sensor is provided as a spare foruse by maintenance personnel, should a failure occur in any of the twosensors.

For airplanes Pre-Mod. SB 145LEG-30-0033, each windshield elementhas a dedicated temperature controller that receives a signal from theassociated temperature sensors and controls the windshieldtemperature. When the temperature reaches the upper limit (45°C),power supply to the heater is interrupted. When the temperature isbelow the lower limit (40°C), power supply is automatically restored. Acaution message W/S HEAT FAIL is presented on the EICAS when asystem failure is detected or the windshield temperature exceeds55°C.

AIRPLANEOPERATIONS

MANUAL



9 Code

01

For airplanes Post-Mod. SB 145LEG-30-0033 or with an equivalentmodification factory-incorporated, the temperature controller has twomodes of operation, defog heat and anti-ice heat mode. When thewindshield heating push button is set to ON, the controller continuouslymonitors the windshield temperature. As temperature drops below26°C (defog mode), it modulates power input to the electric conductivegrid and maintains this temperature. If ice detectors sense iceformation, the controller automatically increase power input to maintainthe temperature at 43°C (anti-ice mode). If both ice detectors areinoperative, the Override knob on the Overhead Panel set to ALLposition provides manual means to put both systems into anti-ice modeautomatically increasing power input to maintain the temperature at43°C. A caution message W/S HEAT FAIL is presented on the EICASwhen a system failure is detected or the windshield temperatureexceeds 65°C.

EICAS MESSAGES


CAUTION

W/S 1 (2) HEAT FAIL For airplanes Pre-Mod. SB145LEG-30-0033, associatedwindshield heating systemfailure (< 38°C) or associatedoverheat condition (> 55°C).For airplanes Post-Mod. SB145LEG-30-0033, associatedwindshield heating systemfailure or associated overheatcondition (> 65°C).

SENSOR HEATING SYSTEMThe Sensor Heating System provides automatic operation for theheater elements of Pitot tubes 1 and 2, Pitot/Static 3, PressurizationSystem and ADS Static Ports, TAT sensors 1 and 2, and AOA vanes 1and 2, thus providing constant temperature and ice-free operationduring all flight phases.

All the sensors are electrically heated and controlled by three buttons,located on the overhead panel.

In the automatic mode, the sensor heating system operates accordingto three functional logic’s:


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MANUAL

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01 REVISION 5

− Pitot 1 and 2 and Pitot/Static 3, AOA 1 and 2, ADS Static Ports 1, 2,3 and 4, and Pressurization Static Ports 1 and 2 are heatedwhenever at least one engine is running (N2 above 54.6%).

− A separated logic assures Pitot/Static 3 and Pressurization SystemStatic Port 2 heating in any flight condition.

− TAT 1 and 2 are heated provided any of Engine 1 or 2 anti-icingsubsystem is functioning or airplane is in flight.

Heater deactivation is accomplished either when the above conditionsare not met or when the associated control button is manually pressed.

Caution messages are presented on the EICAS to indicate that thesensor heating is inoperative. These messages are inhibited during thetakeoff and approach phases.

EICAS MESSAGES

TYPE MESSAGE MEANINGPITOT 1 (2, 3) INOP − Associated sensor heating

inoperative with any enginerunning (N2 above 60%).

− Both engines N2 below 50%.

CAUTIONAOA 1 (2) HEAT INOP − Associated sensor heating

inoperative with any enginerunning (N2 above 60%) andairplane airborne.

− Both engines N2 below 50%.

TAT 1 (2) HEAT INOP Associated sensor heatinginoperative in icing conditionsand airplane airborne.

LAVATORY WATER DRAIN AND NIPPLEHEATING SYSTEMThe lavatory waste water drain and potable water service nipples(overflow and fill) are heated by electric resistors to prevent clogging bywater freezing under any atmospheric conditions on the ground and inflight.

The heating is automatically turned on when the DC BUS 1 is powered.Refer to Section 2-2 – Equipment and Furnishings.

AIRPLANEOPERATIONS

MANUAL



11 Code

01

ICE PROTECTION CONTROLS AND INDICATORS

ICE PROTECTION CONTROL PANEL

1 - ENGINE AIR INLET ANTI-ICING BUTTONS− Turns off (released) or permits (pressed) the automatic

activation of the associated engine air inlet anti-icing subsystem.− A striped bar illuminates inside the button to indicate that it is

released.− An OPEN inscription illuminates inside the button to indicate that

the associated engine air inlet anti-icing valve is open.

2 - WING ANTI-ICING BUTTON− Turns off (released) or selects the automatic mode (pressed) of

the half-wing anti-icing subsystems.− A striped bar illuminates inside the button to indicate that it is

released.− An OPEN inscription illuminates inside the button to indicate the

following conditions:− Both valves are open with the system commanded to open.− At least one valve is open with the system not commanded to

open.

3 - HORIZONTAL STABILIZER ANTI-ICING BUTTON− Turns off (released) or permits (pressed) the automatic

activation of the horizontal stabilizer anti-icing subsystem.− A striped bar illuminates inside the button to indicate that it is

released.− An OPEN inscription illuminates inside the button to indicate that

the horizontal stabilizer anti-icing valve is open.

4 - SENSOR HEATING BUTTONS− The left button controls Pitot tube 1, AOA 1 vane, TAT 1 probe,

ADC Static Ports 1 and 3, and pressurization static port 1.− The central button controls Pitot/Static tube 3 and pressurization

static port 2.− The right button controls the Pitot tube 2, AOA 2 vane, TAT 2

probe and ADC static ports 2 and 4.− When pressed, the associated sensor heating system operates

in the automatic mode according to its functional logic. Whenreleased, the associated sensor heating system is turned off.

− A striped bar illuminates inside the button to indicate that it isreleased.


AIRPLANEOPERATIONS

MANUAL

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01 REVISION 5

5 - ICE DETECTION TEST KNOB

Permits the half-wing, horizontal stabilizer and engine air inlet anti-icing subsystems to operate for test purposes, by simulating anicing condition on ice detectors 1 and 2. The adequate systemoperation is confirmed by the illumination of the OPEN inscriptionsin the anti-icing buttons, which indicate the current valve position.

NOTE: The ICE CONDITION, ICE DET 1 (2) FAIL and BLD 1 (2)LOW TEMP messages are displayed during test. TheCLR ICE 1, CLR ICE 2, CLR/I INOP 1 and CLR/I INOP 2caution messages are displayed on the EICAS and thelights CLR ICE 1 and CLR ICE 2 illuminate only when theIce Detection Test Knob is selected to 1.

6 - ICE DETECTION OVERRIDE KNOB

ENG - Turns on the engine air inlet anti-icing subsystems forground speeds below 25 knots. Above 25 knots the wingand horizontal stabilizer anti-icing subsystems are alsoturned on if icing condition is detected.

AUTO- Allows the automatic operation of the bleed air anti-icingsystem.

NOTE: If ground speed is equal or above 25 knots and anicing condition is detected, wing and horizontalstabilizer anti-icing subsystems are turned on. Theengine anti-icing subsystem is turned on as soonas an icing condition is detected.

ALL - Turns on the complete bleed air anti-icing system providedairplane is in flight.

NOTE: On ground, below 25 knots, only engine anti-icing isturned on.

7 - WINDSHIELD HEATING BUTTON

− Turns on (pressed) or turns off (released) the windshield heatingsystem.


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ICE PROTECTION CONTROL PANEL


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ICE DETECTION SYSTEMIce detectors 1 and 2 are respectively installed at the airplane’s left andright nose section, to provide icing condition detection.

The ice detector was designed to pick up ice quickly. Therefore, in themost cases, ice will be detected before it can be noticed by the crew.

NOTE: Notwithstanding ice detector monitoring, the crew remainsresponsible for monitoring icing conditions and for manualactivation of the ice protection system if icing conditions arepresent and the ice detection system is not activating the iceprotection system.

A 0.5 mm (0.020 inch) ice thickness, on any probe, causes bleed airanti-icing system automatic mode activation, a SPS angle of attack setvalues reduction (refer to Stall Protection System on Section2-4 – Crew Awareness), and an advisory message to be presented onthe EICAS.

During ice encounters, the icing signal remains active during 60seconds. Simultaneously, an internal ice detector heater is activated tode-ice the unit and probe. When the probe’s natural frequency isrecovered, heating is de-energized. Once deiced, the sensing probecools within a few seconds and is ready to once more monitor icebuild-up. Then a new detection cycle begins and remains as long asthe ice condition persists.

In case of failure of any or both ice detectors, a caution message ispresented on the EICAS and the bleed air thermal anti-icing system maybe activated through the OVERRIDE knob on the Ice Detection panel.The system’s normal operation may be checked through the TESTknob on the Ice Protection panel.

WING CLEAR ICE DETECTION SYSTEM

The clear ice phenomena may occur on the wing upper surfaces whenthe airplane performs a prolonged operation in high altitude with thewing fuel tanks quantity is kept above 70% until ground. Then, if theairplane is exposed to conditions of high humidity, rain, drizzle, or fogeven at ambient temperatures above freezing, the water contained inthe atmosphere, when in contact with the cold wing, may condenseand freeze.

Once the clear ice accumulation is difficult to be detected visually,clear ice detectors 1 and 2 are respectively installed on the airplane’sleft and right wing upper surface.


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MANUAL

2-15-15 Page

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01 REVISION 5

Each sensor is flush mounted with the surface and consists of avibrating diaphragm. Any ice build-up is sensed by the diaphragm anda caution message is presented on the EICAS and the Clear IceIndication Light illuminates.

In case of failure of any or both ice detectors, a caution message is alsopresented on the EICAS.

The system’s normal operation may be checked through the TESTknob on the Ice Protection panel.

EICAS MESSAGES

TYPE MESSAGE MEANINGICE DETECTORS FAIL Both ice detectors have

failed.ICE DET 1 (2) FAIL Associated ice detector

has failed.CAUTION CLR ICE 1 (2) Ice build-up over the left

or right wing uppersurface.

CLR/I INOP1 (2) Associated clear icedetector has failed.

ADVISORYICE CONDITION Airplane is flying under

icing conditions.

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MAIN PANEL

1- CLEAR ICE LIGHT

− Color: amber− Illuminates CLR ICE 1 to indicate that there is ice build-up over the

left wing upper surface.− Illuminates CLR ICE 2 to indicate that there is ice build-up over the

right wing upper surface.


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01 REVISION 5

WINDSHIELD WIPER SYSTEM (IF INSTALLED)A two-speed windshield wiper is provided for the left and rightwindshields. Each system comprises a motor-converter, a wiper arm,and blades. A control box provides speed control, synchronization, andoff-screen park functions for both systems through independentchannels.

Each system has its own independent power supply and a four-positionknob on the overhead panel.

WINDSHIELD WIPER CONTROL PANEL

1 - WINDSHIELD WIPER SELECTOR KNOB

TIMER - Provides intermittent operation of the associatedwindshield wiper in single cycles (two strokes) with an 8second time interval between two cycles, in high speed.

OFF - Associated wiper blades travel to the windshield inboardposition, parking out of pilots vision.

LOW - Associated wiper operates at approximately 80 strokes perminute.

HIGH - Associated wiper operates at approximately 140 strokesper minute.

NOTE: Dry windshield operation leads the motor-converter to a stallcondition, due to the high friction level. The controllersenses the motor-converter current surge and drives thearm directly to the parked position. The system remainsinoperative until the Windshield Wiper Selector Knob is setto OFF position and a new operation mode is selected.

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01

WINDSHIELD WIPER CONTROL PANEL


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01 REVISION 5


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NAVIGATION ANDCOMMUNICATION


1 Code

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SECTION 2-18

NAVIGATION AND COMMUNICATION

TABLE OF CONTENTS

Block Page

General ...............................................................................2-18-01 .. 01

Radio Management System (RMS) ....................................2-18-05 .. 01Integrated Communication Unit (RCZ-851E) ..................2-18-07 .. 01Integrated Navigation Unit (RNZ-851).............................2-18-09 .. 01Radio Management Unit (RMU)......................................2-18-11 .. 01

RMU Pages .................................................................2-18-11 .. 01RMU Normal Operation ...............................................2-18-11 .. 03RMU Abnormal Operation ...........................................2-18-11 .. 09RMU Controls and Indicators ......................................2-18-11 .. 10

Tuning Backup Control Head ..........................................2-18-13 .. 01Normal Mode ...............................................................2-18-13 .. 01Emergency Mode ........................................................2-18-13 .. 01Self-Test ......................................................................2-18-13 .. 01TBCH Controls and Indicators.....................................2-18-13 .. 02

Digital Audio Panel ..........................................................2-18-15 .. 01Normal Mode ...............................................................2-18-15 .. 01Emergency Mode ........................................................2-18-15 .. 01Digital Audio Panel Controls and Indicators ................2-18-15 .. 03

Communication Controls and Indicators .........................2-18-20 .. 01

HF Communication System - KHF-950 (∗) .........................2-18-21 .. 01HF Operating Modes.......................................................2-18-21 .. 01HF Normal Operation KCU-951 CONTROL UNIT ..........2-18-21 .. 03HF Controls and Indicators .............................................2-18-21 .. 08HF Normal Operation KFS-954 CONTROL UNIT...........2-18-21 .. 13HF Controls and Indicators .............................................2-18-21 .. 19

Third VHF Communication System - Collins 22A (∗)..........2-18-22 .. 01Third VHF COM Controls and Indicators ........................2-18-22 .. 01

Third VHF Communication System - HoneywellRS-833/853(∗) ....................................................................2-18-22 .. 07

Third VHF COM Controls and Indicators ........................2-18-22 .. 12

NOTE: Optional equipment are marked with an asterisk (∗) and itsdescription may not be present in this manual.


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SELCAL System (∗) ...........................................................2-18-23.. 01SELCAL Controls and Indicators ....................................2-18-23.. 02

Honeywell Mark III CMU (∗)................................................2-18-24.. 01CMU Normal Operation ..................................................2-18-24.. 04CMU Abnormal Operation ..............................................2-18-24.. 04CMU Controls and Indicators..........................................2-18-24.. 06Printer Controls and Indicators .......................................2-18-24.. 08

Aircraft Communication Addressing andReporting System (ACARS) (∗) ............................2-18-24.. 11

ACARS Operation...........................................................2-18-24.. 14ACARS Controls and Indicators .....................................2-18-24.. 15ACARS Printer Controls and Indicators..........................2-18-24.. 18

Cockpit Voice Recorder......................................................2-18-25.. 01Self-Test .........................................................................2-18-25.. 01Erase Function................................................................2-18-25.. 01Cockpit Voice Recorder Controls and Indicators............2-18-25.. 02

Passenger Address System ...............................................2-18-27.. 01Passenger Address Operating Modes............................2-18-27.. 02

Satcom System (∗) .............................................................2-18-28.. 01Introduction .....................................................................2-18-28.. 01Satcom Operation...........................................................2-18-28.. 01Satcom Controls and Indicators......................................2-18-28.. 05

Iridium Stellite Telecommunication System (∗) ..................2-18-29.. 01Iridium Controls and Indicators .......................................2-18-29.. 02Iridium Operation ............................................................2-18-29.. 04

Inertial Reference System (IRS) (∗) ...................................2-18-30.. 01Inertial Reference System Components .........................2-18-30.. 04IRS Operating Modes .....................................................2-18-30.. 05IRS Operating Procedures..............................................2-18-30.. 10IRS EICAS Messages.....................................................2-18-30.. 12IRS Controls and Indicators............................................2-18-30.. 14IRS Indications on the PFD ............................................2-18-30.. 16

Flight Management System (FMZ 2000) (∗).......................2-18-35.. 01FMS Operating Modes....................................................2-18-35.. 04FMS Controls and Indicators ..........................................2-18-35.. 07

Navigation Displays ............................................................2-18-40.. 01Displays Controls and Indicators ....................................2-18-40.. 02


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Weather Radar System...................................................... 2-18-45 ..01General........................................................................... 2-18-45 ..03Weather Radar Normal Operation ................................. 2-18-45 ..04Interpreting Weather Radar Images............................... 2-18-45 ..04Radar Warm Up Period.................................................. 2-18-45 ..06Ground Operation Precautions....................................... 2-18-45 ..06Weather Radar Operating Modes and Functions........... 2-18-45 ..07Radome.......................................................................... 2-18-45 ..18Weather Radar Controls and Indicators......................... 2-18-45 ..19

Lightning Sensor System (LSS) (∗).................................... 2-18-50 ..01LSS Operation ................................................................ 2-18-50 ..02LSS Controls and Indicators........................................... 2-18-50 ..05

Identification Friend or Foe System (IFF) (∗) ..................... 2-18-80 ..01Selector Panel ................................................................ 2-18-80 ..02IFF Transponder Controls and Indicators....................... 2-18-80 ..04

Precision Area Navigation (P-RNAV) (*) ............................2-18-85...01Limitations.......................................................................2-18-85...01P-RNAV System .............................................................2-18-85...03Normal Procedures.........................................................2-18-85 ...04Contingency Procedures.................................................2-18-85...06Incident Reporting...........................................................2-18-85...07



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01 ORIGINAL


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GENERALThe standard EMB-135 BJ navigation and communication resourcesare provided by the Radio Management System (RMS). The RMS iscontrolled through two Radio Management Units (RMU 1 and 2), anauxiliary control unit, the Tuning Backup Control Head (TBCH), andthree individual Digital Audio Panels (DAP).

The two RMUs provide radio frequency and mode control.Alternatively, the RMU 2 frequencies may be selected through theTBCH.

The Audio System is controlled via three individual Digital AudioPanels, available for the captain, copilot and observer.

The Radio Management System also provides interface with thePassenger Address System, Aural Warning Unit and Cockpit VoiceRecorder.

Optional communication equipment includes an HF transceiver, ThirdVHF NAV/COM, SELCAL and Aircraft Communication Addressing andReporting System (ACARS).

The navigation may be performed using only the standard navigationradio sensors, or using the Flight Management System (FMS)resources. The FMS is an optional equipment that uses the standardnavigation radio sensors, GPS (Global Positioning System) sensors,and, also optionally, the IRS (Inertial Reference System) for positioningand navigation.

Heading inputs to the Integrated Navigation Unit are provided by theAHRS (Attitude and Heading Reference System) or by the IRS. Theseequipment also provide roll and pitch attitudes for the ElectronicAttitude Director Indicator (EADI).

The navigation information is normally presented on the PFD and MFDand may also be available on the RMU, through its navigation backuppage.


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RADIO MANAGEMENT SYSTEM (RMS)

The EMB-135 BJ models are equipped with a Radio ManagementSystem (RMS) that provides management of the following equipmentand associated functions:

− Dual VHF COM;− Dual VHF NAV (VOR, LOC, GS and Marker Beacon);− Single or dual (optional) ADF;− Single or dual (optional) Transponder (ATC and Mode S);− TCAS;− MLS (optional);− Single or dual (optional) DME (including DME Hold);− Digital Audio Panel.

The RMS consists basically of the following major components:

− Remote mounted:− Integrated Navigation Unit (INU);− Integrated Communication Unit (ICU).

− Cockpit Mounted:− 2 Radio Management Unit (RMU);− 1 Tuning Backup Control Head (TBCH);− 3 Digital Audio Panel (DAP).

With the exception of the Digital Audio Panel, all components of theRMS are connected through the digital Radio System Buses (RSB)that allows complete control and information exchange between theunits of the entire RMS. Audio switching control is provided by meansof the controls on the Digital Audio Panel itself. The audio signals aretransmitted from the remote units to the Digital Audio Panel throughdedicated digital audio buses.

The navigation and communication data are displayed on the RMU,PFD and MFD displays.


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RMS SCHEMATIC

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INTEGRATED COMMUNICATION UNIT (RCZ-851E)

The Integrated Communication Unit incorporates an internal VHFcommunication transceiver module and the ATC transponder modulewhich interfaces through a cluster module to the Radio System Bus foroperation.

This unit provides digitized audio signals to the Digital Audio Panel andconventional analog audio interfaces to other systems. The followingmodules are provided in this unit:

− VHF Communication Transceiver Module (TR-850) - This module isa conventional VHF COM transceiver that operates in the frequencyrange of 118 to 136.975 MHz.

− ATC Mode S Transponder Module (XS-850) - The Mode STransponder module has the encoding and decoding capabilityrequired for Mode S operation in addition to the capability to operateas a conventional Air Traffic Control Radio Beacon Service(ATCRBS) transponder. The Mode S operation allows digitaladdressing of an individual airplane and the transmission ofmessages back and forth between the air and the ground.

− ATCRBS Transponder Module (XS-850A) - This transponder moduleprovides only conventional ATC Radio Beacon System transpondercapabilities.

− Mode S Diversity Transponder Module (XS-852) - This transpondermodule provides full ATCRBS, Mode S and TCAS datacommunications capability.

− TCAS Interface Module (XI-851) - The Interface Module allows theIntegrated Communication Unit (ICU) to interface with separateMode S diversity transponder and TCAS. The TCAS interfacemodule replaces the XI-851 Mode S transponder module wheninstalled in the ICU.


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INTEGRATED NAVIGATION UNIT (RNZ-851)

The Integrated Navigation Unit is a complete self-contained navigationsystem. The system consists of the VOR, localizer, glide slope andmarker beacon receiver modules, the ADF module, a six-channelscanning DME module, and audio digitizers. The system alsoincorporates two L-Band antenna (optional), two ADF antenna(optional), two MB antenna, two VOR/ILS antenna and one GS dualantenna.

The following modules are provided in this unit:

− VHF NAV Receiver Module (NV-850) - The VHF NAV receiver is amodule of the Integrated Navigation Unit and houses the majornavigation functions of the VOR/LOC receiver, glide slope receiverand marker beacon receiver.The ILS meets Category II instrument landing requirements.Housed within the NAV receiver is a glideslope receiver whichprovides 40 channels of glideslope information for the conventionalILS. Also includes a 75 MHz marker beacon receiver which detectsand transmits the tones of the marker beacons to the Audio System.

− DME Transceiver Module (DM-850) - The DME module is asix-channel DME that simultaneously tracks four selected channelsfor distance, groundspeed and time to station as well as monitoringtwo additional channels for the ident functions. This feature gives thesystem the capability of tracking four channels and having thedecoded identifier readily available from two additional channels. Theunit dedicates two of the four selected channels to the FMS (ifinstalled). Thus, with the FMS installed, there are two remainingchannels to control and display ident, distance, time to station andground speed. Even with the FMS installed, the preset or standbyVOR channel, when selected, provides instant station identificationsince it was one of the two additional channel being monitored.

− ADF Receiver Module - The ADF System comprises the ADFreceiver (DF-850) and the companion ADF antenna (AT-860). TheADF receiver operates in the frequency ranges of 100 to 1799.5 kHzand 2181 to 2183 kHz (marine emergency frequency range).


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RADIO MANAGEMENT UNIT (RMU)

The Radio Management Unit consists of a display and a bezel panelthat provide control of the communications and radio navigationequipment. Additional airplane systems information is also available onspecific RMU selectable pages.

The EMB-135 BJ is equipped with two RMUs, each one responsible forcontrolling the on-side radio equipment (e.g., RMU 1 controls theNAV/COM 1). However, through the cross-side operating mode it ispossible to select the opposite side radio frequencies.

There is no master switch for the RMUs: when the airplane isenergized, both RMUs (and the EICAS) are automatically turned ON.However, only the COM 1 radio is available (dashes on the remainingRMUs fields) until the AVIONICS MASTER is switched ON.

Additionally, in the event of an electrical emergency the RMU is abackup display for the main panel (PFDs and MFDs). In this conditionthe main panel is turned off and the NAVIGATION Backup Page, thatpresents basic navigation information, may be accessed through RMUpage.

RMU PAGES

Available RMU pages are as follows: RADIO Page, NAV and COMMEMORY Pages, ATC/TCAS Control Page, NAVIGATION BackupPage, ENGINE Backup Pages 1 and 2, SYS SELECT Page (COMband options) and MAINTENANCE Page.

Pressing the Page Control Button (PGE) selects the Page Menu.Pressing the Line Select Button associated with the desired page willcause the respective page to be displayed. The RADIO Page will bedisplayed again when the Line Select Button associated with theRETURN TO RADIOS label is pressed.

RADIO PAGE

Normally presented after power up, the RADIO Page is divided into fivededicated windows. Each window groups the data associated with aparticular function: COM, NAV, ATC/TCAS, ADF and MLS (optional).In addition the windows provide complete control of the frequency andoperating modes of the associated function.


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01 ORIGINAL

NAVIGATION/COMMUNICATION MEMORY PAGES

The Memory Page presents two similar displays called First MemoryPage and Second Memory Page. The First Memory Page showsmemory locations 1 through 6 and the Second Memory Page showsmemory location 7 through 12. Both the COM and NAV Memory Pagesare functionally identical.

ATC/TCAS CONTROL PAGE

The ATC/TCAS Control Page allows the pilot to select various TCASoperational features:

• Intruder Altitude− REL: Target’s altitude displayed relative to one’s own airplane

(default).− FL: Target’s altitude displayed as flight level (reverts to REL

after 20 sec).• TA Display

− AUTO: Traffic targets displayed only when TA or RA targetconditions exist.

− MANUAL: All traffic targets displayed within the viewingairspace.

• Flight Level ID Allows Mode S coding to reflect the current flight’s call sign.• Flight Level 1/2

Display of the transponder’s encoded altitude and the air datasource for that altitude.

NAVIGATION BACKUP PAGE

The NAVIGATION Backup Page consists of a backup navigationdisplay that presents HSI, MB, DME, NAV (VOR) and ADF information.

ENGINE BACKUP PAGE

The ENGINE Backup Page displays information normally presented onthe EICAS, as engine and systems indications, as well as EICASmessages. The ENGINE Backup Page is divided into two pages. Thefirst presents only engine indications, while the second presentssystems indications and EICAS messages. For further information onEngine Backup Page refer to Section 2-10 - Powerplant and 2-4 - CrewAwareness.

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SYSTEM SELECT PAGE

The SYS SELECT Page allows the selection of COM 1 and COM 2between Narrow and Wide bands.

MAINTENANCE PAGE

This page displays test results information depending upon the type oftest that is being carried out (power on self-test or pilot activated self-test). Two pages may be presented if a failure is detected, depending ifthe failure is in the RMS or in one of the radios. This page is notavailable in flight.

RMU NORMAL OPERATION

RMU SELF-TEST

On the ground, the RMS performs a self-test each time power isapplied after power off periods greater than 10 seconds. This testmonitors the primary and secondary radio buses as well as theindividual Radio Systems for proper operation. Each function teststatus is displayed in its respective window.

Under normal conditions, the COM will be operational within 7 secondsafter power on and the remaining radio equipment units within 50seconds. The test can be terminated by pressing the Test Button in theRMU Bezel Panel.

If any bus or radio test parameter failure occurs, an associated errormessage will be displayed on the test failure window, below the COMand NAV windows. Radio System failures are displayed in the firstfailure window and function failures in a second failure window. Thefailure windows may be removed by pressing and holding the TestButton. If the test is successfully completed the RMU will display theRADIO Page with the same radio configuration prior to the last powerdown.

NOTE: Any radio equipment that is not powered up when the test isinitiated by the RMU will generate an error message.


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Additionally the pilot may perform a test by pressing the Test Button onthe RMU Bezel Panel which causes the activation of the self-test of thecomponent associated with the window in which the yellow cursor islocated. Upon successful completion of this test, a PASS message willbe displayed for a short time in the window, indicating the successfulcompletion of the test. If this test is not successful completed, an errormessage (ERR) will be displayed in the window.

NOTE: Errors detected by the self-test indicate one or more parameteroutside their self-test limit but may not necessarily indicate non-operation of the function. The pilot should verify the operationof the function.

CROSS-SIDE OPERATION

The RMU is provided with a feature called cross-side operating mode.This feature allows the RMU to be changed from its normal operatingmode of tuning the on-side radio equipment to the mode of tuning theopposite side radio equipment.

The cross-side operation is selected by pressing the cross-sideTransfer Button, labeled 1/2, on the RMU Bezel Panel, with the yellowcursor box in any window, except the ATC/TCAS window. The entireRMU display and operation is transferred from the opposite side to theside that has commanded the Cross-side Operating Mode. If the yellowcursor box is in the ATC/TCAS window, pressing the cross-sideTransfer Button selects which transponder (1 or 2) will be in operation.

In the cross-side operation, the RMU Window/Control Side Ident will bedisplayed in magenta on the side that has selected the operation andany change made will be displayed in yellow on the opposite side RMUto indicate that the change was carried out remotely.

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COM OPERATION

The normal COM operation is enabled with the RMU RADIO Pagedisplayed. The COM window has two frequency lines. The upper linedisplays the active COM frequency while the lower line displays thepreset frequency. Pressing the Line Select Button associated with thepreset frequency will cause the yellow cursor box to move to enclosethat frequency. In this condition the enclosed preset frequency may bechanged through the Frequency Tuning Knobs. When the FrequencyTuning Knobs are actuated the label MEMORY and the associatedmemory location number, both below the lower frequency line, willchange to a TEMP label indicating that the new preset frequency is notyet stored in the memory of the RMU. Frequency storage may beaccomplished by pressing the Memory Storage Button, labeled STO,on the RMU Bezel Panel. This action will also provide the previousMEMORY label and the associated memory location number toreplace the TEMP label, indicating that the new preset frequency hasbeen stored in the indicated memory location.

Placing the yellow cursor box to enclose the MEMORY label, bypressing a second time the Line Select Button beside the COMwindow, will allow scrolling through the entire RMU stored memory.This may be performed by rotating the Frequency Tuning Knob eitherclockwise to memory location increment or counterclockwise todecrement.

The exchange between the active frequency displayed in the upper lineof the window and the preset frequency displayed in the lower line maybe accomplished by pressing the Frequency Transfer Button on theupper left corner of the RMU Bezel Panel. This effectively causes theCOM to change to the new active frequency that previously was thepreset frequency. In this condition, the previous active frequency dropsdown to the second line of the COM window and becomes a temporarypreset frequency. This is indicated by the TEMP label displayed underthat frequency. The TEMP label also indicates, in this case, that thefrequency displayed in the second line has not been stored in amemory location.

NOTE: The RMU controls the third VHF for airplanes equipped withHoneywell Third VHF System RCZ-833/853 models.


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• Direct COM Tuning

Direct COM tuning is accomplished by pressing and holding forapproximately 3 seconds the Line Select Button beside the COMpreset frequency line. The yellow cursor box will enclose the activefrequency allowing direct COM tuning to that frequency, and the presetfrequency line will be blank.

To exit from direct COM tuning, press and hold the Line Select Buttonbeside the preset frequency line, until the preset frequency appears onthe COM window.

• Squelch Function

The COM squelch function is controlled through the Squelch ControlButton, labeled SQ, on the RMU control bezel. Pressing this button willcause the COM radio to open its squelch and allow any noise or signalpresent in the receiver to be heard in the Audio System. The squelchopen condition is indicated by the SQ label displayed on the top of theCOM window. Pressing the Squelch Control Button again will close theradio squelch immediately.

• Automatic Time-Out

After approximately two minutes of continuous transmission, thetransceiver turns its transmitter off and a beep sound in the audio systemalerts the pilot to the fact. The transceiver then reverts to receiver modein order to prevent a stuck microphone button from blocking thecommunications channel. Should the time-out occur, the pilot can reset itby simply releasing the push to talk button and pressing it again.

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NAV OPERATION

The NAV operation is identical to the COM operation. However, NAVcontrols are accomplished by actuation of the Frequency TransferButton and the Line Select Button located on the upper RH of the RMUBezel Panel. Furthermore, the NAV window has an additional functioncalled DME Split Tuning Mode. The operation in the DME Split TuningMode is similar to the operation in the DME Hold Mode.

The NAV system also incorporates FMS autotuning capability. Throughthe NAV Memory Page it is possible for the FMS to perform automatictuning of the navigation radios (raw data) along the route by pressingthe upper RH Frequency Transfer Button, which enables or disablesthe FMS autotuning capability. When the VOR or the ILS frequency isautotuned by the FMS, a magenta VOR or ILS frequency and amagenta AUTO label will be displayed on the top border of the RADIOPage NAV window.

DME OPERATION

In the normal DME operations only one of the six DME channels ispaired with the VOR active frequency and one other with the presetVOR frequency. However, pressing the DME Select Button, labeledDME, on the RMU Bezel Panel, will enable the DME to be tunedindependently of the VOR active frequency.

Pressing the DME Select Button once will cause the NAV window tosplit into two windows. The top window will display the active VORfrequency and the lower window, with the DME label, will display theactive DME frequency in VHF format. When the NAV window is split,an H (DME Hold) label is displayed in the DME window to indicate thatthe DME is not paired with the active VOR/ILS frequency. In this casethe DME hold condition will also be announced on the PFD. In thiscondition, the DME may be tuned directly by simply pressing theassociated Line Select Button beside the DME window and tuning thenew DME channel through the Frequency Tuning Knobs.

Pressing the DME Select Button again will cause the frequency to bedisplayed in the channel format (TACAN).

Pressing the DME Select Button for the third time will cause the NAVwindow to resume its normal mode with the active and presetfrequencies being displayed while returning the DME to the condition ofchanneling with the active VOR frequency.


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ADF OPERATION

The tuning of ADF frequencies is similar to that performed on theairplane’s other radios equipment. Pressing the Line Select Buttonbeside the ADF frequency display will move the yellow cursor box tosurround the ADF frequency in the RMU display. Then, slowly turningthe Frequency Tuning Inner Knob clockwise causes the ADFfrequencies to advance in 0.5 kHz increments while slowly turning theouter knob clockwise will cause the frequencies to advance in 10 kHzincrements. ADF tuning through the Frequency Tuning Knobs isaccomplished using proportional rate. If the knobs are turned in slowdeliberate steps the frequency will follow likewise. However, if the knobis turned rapidly, the frequency will skip several steps, depending uponthe speed at which the knob is turned. This allows accomplishing largefrequency changes with a very slight rotation of the knob.

The RMU also has the capability of storing an ADF frequency. This isaccomplished by selecting the desired ADF frequency and thenpressing the Memory Storage Button on the RMU Bezel Panel. Toretrieve the stored frequency from memory, the ADF frequency LineSelect Button must be pressed for 2 seconds.

The ADF is provided with a mode control capability. ADF operationalmodes can be selected by moving the yellow cursor box to the ADFmodes field in the ADF window and then pressing the Line SelectButton beside the ADF modes field or rotating the Frequency TuningKnobs. Repeatedly pressing the Line Select Button will cause themodes to step in one direction while rotating the Frequency TuningKnobs will select the modes either up or down the current location.

The ADF operational modes are the following:

- ANT - The ADF receives signal only.

- ADF - The ADF receives signal and calculates relative bearings tostation.

- BFO - The ADF adds a beat frequency oscillator for reception ofCW signals.

- VOICE - The ADF opens width of IF bandwidth for better auralreception.

NOTE: Bearing information is available in the ADF and BFO modesonly.

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TRANSPONDER AND TCAS OPERATION

Transponder operation is similar to other radio equipment since itrequires moving the yellow cursor box to a desired function. In order totune a desired ATC code, press the Line Select beside the ATC codedisplay. This action will enable the Frequency Tuning Knobs to changethe ATC codes. The outer knob sets the thousands and hundredsdigits and the inner knob sets the tens and ones digits.

Pressing and holding the code Line Select Button will recall the storedpreset code (typically used for VFR). A new code may be stored bysetting the code and then pressing the Memory Storage Button on theRMU Bezel Panel.

Pressing the Line Select Button associated with the transponderoperating mode display will move the yellow cursor box to surround themode annunciation in the ATC/TCAS window allowing to set a newtransponder mode if a non-standby mode is selected. Once the modeannunciation is surrounded, pressing the Transfer Button 1/2 will selectwhich transponder will be in operation (e.g., 1 ATC ON to 2 ATC ON).

The transponder operational modes are the following:

− ATC ON - Replies on Modes S and A, no altitude reporting.− ATC ALT - Replies on Modes A, C and S, with altitude reporting.− TA ONLY - TCAS Advisory Mode is selected.− TA/RA - TCAS Traffic Advisory/Resolution Advisory Mode is selected.

RMU ABNORMAL OPERATION

Loss of the Primary Radio Bus will disable the cross-side controlcapability and also the TBCH. However, no radio functions will be lost.The radios on both sides will still be functional through the SecondaryRadio Buses.

Loss of the left and/or right Secondary Radio Bus will not disable theradio functions. The radios may be tuned, in this condition, through thePrimary Radio Bus or through the cross-side control feature.

As a safety feature of the RMU, if any component of the Radio Systemfails to respond to the commands from the RMU, the frequencies orthe operating commands associated with that particular function will beremoved from the RMU display and replaced with dashes.


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RMU CONTROLS AND INDICATORS

RMU BEZEL PANEL

1 - FREQUENCY TRANSFER BUTTON− When pressed, the active frequency (upper line) and the preset

frequency (lower line) in the COM or NAV windows exchangelocation and function.

2 - LINE SELECT BUTTONS− The first press of the button moves a yellow cursor box to

surround the data field associated with that particular Line SelectButton. This enables the Frequency Tuning Knobs to change thedata or the mode marked by the cursor. For some functions,additional pressing of the Line Select Button will toggle modes orrecall stored frequencies. The Line Select Buttons, if keptpressed, allows ADF and ATC memories to be recalled, and toenter or exit Direct Tune Mode for COM and NAV.

3 - FREQUENCY TUNING OUTER KNOB − Allows the data field enclosed by the cursor to be modified. The

data may be frequency setting, stored frequencies or mode,depending upon the data field. When setting a frequency, thisknob controls the digits to the left of the decimal point.Furthermore, this knob also controls the RMU brightness, whichis enabled by pressing the Dimming Button.

4 - FREQUENCY TUNING INNER KNOB − Is functionally similar to the Frequency Tuning Outer Knob

except that when setting the frequency, this knob controls thedigits to the right of the decimal point.

5 - MEMORY STORAGE BUTTON− Pressing this button will cause a temporary (TEMP) COM or

NAV pre-select frequency to be stored in the memory andassigned numbered location, provided the cursor has first beenplaced around that frequency.

NOTE: ADF and ATC have only one memory location.

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11 Code

01

6 - DME SELECT BUTTON− Allows selection of the DME Hold Mode, tuning a different DME

channel, not paired with the VOR/ILS frequency, without changingthe active VOR frequency. Repeated pressing of this buttonenables display and selection of the DME channels in VHF andTACAN formats, and then back to the paired VOR/DME mode.

7 - CROSS-SIDE TRANSFER BUTTON− With the cursor in any window, except the ATC or TCAS display,

pressing this button will transfer the entire RMU operation anddisplay from the cross-side system.

− With the cursor in the ATC or TCAS window, pressing thisbutton selects which transponder will be in operation.

− With enhanced TCAS, the button allows control of TCAS data inthe cross-side display.

8 - TEST BUTTON− When pressed, causes the component associated with the

present position of the yellow cursor box to activate its internalself-test circuits for a complete end-to-end test of the function.To properly accomplish the equipment self-test, the Test Buttonmust be pressed and held down as follows:− About 2 seconds for COM transceiver self-test.− From 5 to 7 seconds for DME, ATC and ADF self-test.− About 20 seconds for NAV (VOR/ILS) self-test.

− Releasing the Test Button at any time immediately returns theequipment to its normal operation in the actual function.

− If the Test Button is held pressed for 30 seconds or more, theradios are automatically commanded back into normal operation.

9 - PAGE CONTROL BUTTON− Provides access to the page menu.

10 - DIMMING BUTTON − The RMU features an automatic screen brightness adjustment,

within a limited range, to keep the display visibility optimized.The Dimming Button enables RMU brightness to be controlledmanually through the Frequency Tuning Outer Knob. Themanual dimming control can be disabled by pressing theDimming Button again or any Line Select Button.


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11 - TRANSPONDER IDENTIFICATION MODE BUTTON − Selects the Transponder Identification Response Mode. The

ident squawk will stop after 18 seconds.

12 - SQUELCH CONTROL BUTTON − Causes the COM radio to open its squelch allowing any noise

or signal present in the radio to be heard in the Audio System.The label SQ is displayed on the top line of the COM windowwhen the squelch is open. When pressed a second time theSquelch Control Button closes the squelch.

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SEPTEMBER 30, 2002 2-18-11 Page

13 Code

01

RMU BEZEL PANEL


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14 Code


RMU DISPLAY

PAGE MENU

1 - PAGE MENU IDENTIFICATION− Indicates that Page MENU is selected.− Color: White.

2 - COM AND NAV MEMORY PAGE LABEL− To access the COM or NAV MEMORY Pages press the Line

Select Button adjacent to the desired page.− Color: Green.

3 - ATC/TCAS PAGE LABEL− To access the ATC/TCAS Page press the Line Select Button

adjacent to this label.− Color: Green.

4 - NAVIGATION PAGE LABEL− To access the NAVIGATION Page press the Line Select Button


5 - ENGINE PAGE LABEL− To access the ENGINE Page press the Line Select Button


6 - SYS SELECT PAGE LABEL− To access the SYS SELECT Page press the Line Select Button


7 - MAINTENANCE PAGE LABEL− To access the MAINTENANCE Page press the Line Select

Button adjacent to this label.− Color: Green.

8 - RETURN TO RADIOS PAGE LABEL− To return to the RADIOS Page press the Line Select Button


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15 Code

01

PAGE MENU


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2-18-11 Page

16 Code


RADIO PAGE

1 - PRESET FREQUENCY MEMORY LOCATION (ONLY FOR NAVAND COM WINDOWS)− Identifies the preset frequency as temporary (TEMP label) or

retrieved from the memory (MEMORY label followed by itsmemory location).

− Colors:− Cyan for on-side operation.− Yellow for cross-side operation.

− When marked by the yellow cursor box, the memory locationlabels and their associated stored frequencies can be scrolledby using the Frequency Tuning Knobs.

2 - COM WINDOW/CONTROL SIDE IDENTIFICATION− Identifies the window and which source equipment (side 1 or 2)

is active in that RMU.− Colors:

− White for on-side source.− Magenta for cross-side source.

3 - VHF COM ACTIVE FREQUENCY− Indicates the active frequency for that window.− Colors:

− White for on-side operation.− Yellow for cross-side operation.

− Digits are replaced by dashes in case of any failure in theassociated source.

4 - VHF COM PRESET FREQUENCY− Indicates the preset frequency.− Colors:

− Cyan for on-side operation.− Yellow for cross-side operation.

NOTE: When DME Hold is not selected, the NAV Window alsopresents a similar preset frequency field.

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17 Code

01

5 - NAV WINDOW/CONTROL SIDE IDENTIFICATION− Identifies the window and which source equipment (side 1 or 2)



6 - VHF NAV ACTIVE FREQUENCY− Indicates the active frequency for that window.− Colors:



7 - DME HOLD MODE ANNUNCIATION− Indicates that the DME is in Hold Mode and the active DME

channel is selected separately from the active VOR/ILSfrequency.

− Color: Yellow.

8 - DME STATION IDENTIFICATION CODE− Displays the digital identification code of the ground station to

which the DME is tuned with.− Color: White.

9 - DME HOLD MODE FREQUENCY− Indicates the active frequency in DME Hold Mode operation, in

VHF (represented) or TACAN formats.− Color: White.

10 - ADF WINDOW/CONTROL SIDE IDENTIFICATION− Identifies the window and which source equipment (side 1 or 2)




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18 Code


11 - ADF ACTIVE FREQUENCY− Indicates the active frequency for that window.− Colors:



12 - ADF MODES FIELD− Displays the ADF modes as selected either through the

second ADF Line Select Button (achieved by repeatedpressing) or through the Frequency Tuning Knobs when theyellow cursor box is located in this field.

− Color: Green.

13 - TRANSPONDER OPERATING MODE ANNUNCIATION− Displays the active transponder operating mode as selected

through the Frequency Tuning Knobs when the yellow cursorbox is located in this field. Pressing the Line Select Buttonbeside this field will alternate between the pre-selectedtransponder mode and the standby mode.

− Color: Green.

14 - ATC CODE− Displays the active ATC code number.− Color: White.

15 - ATC/TCAS WINDOW− Identifies the window as the ATC/TCAS window.− Colors: White.

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19 Code

01

RMU RADIO PAGE


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20 Code


COM MEMORY PAGE

1 - MEMORY PAGE IDENTIFICATION− Identifies the page as a COM Memory Page.− Color: White.

2 - ACTIVE COM FREQUENCY− Identifies the COM frequency that is currently active.− Color: White.

3 - SQUELCH MODE INDICATION− Indicates if squelch is open.− Color: Yellow

4 - MEMORY PAGE SELECTED ANNUNCIATION− Indicates that the Memory Page is selected.− Color: Green.

5 - MEMORIES DISPLAY− Displays the preset frequencies and their associated locations.− When there is no frequency stored in a memory location only

the location number will be displayed in the associated memorydisplay line.

− Colors:− Memory identifications are green.− Frequency is cyan.

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21 Code

01

6 - MEMORY INSERT PROMPT− If it is desirable to insert a new frequency in a particular memory

location, simply press the Line Select Button beside the locationline, moving the yellow cursor box to that line. Then press theLine Select Button beside the Insert prompt label. This willcause all the data in memory from the insert location downwardto shift one position down. The cursor will remain in the insertionselected location allowing the new frequency to be tuned andstored in that memory location. A MEM FULL (Memory Full)annunciation will be displayed in the RMU display if the 12memory locations are filled and the Line Select Buttonassociated with the Insert prompt is pressed.

− Color: Green.

7 - MEMORY DELETE PROMPT− To delete a frequency from the memory, press the Line Select

Button adjacent to the line associated with the frequency to bedeleted. Then press the Line Select Button adjacent to theDelete prompt. The frequency enclosed by the cursor will bedeleted from the memory. Higher numbered memory locationswill then move upward to fill the empty memory location.

− Color: Green.

8 - RADIO PAGE RETURN PROMPT− Pressing the associated Line Select Button will return the RMU

display to the RADIO Page.− Color: Green.

9 - MEMORY MORE PROMPT− The More prompt allows to display memory locations 7 through

12, by pressing the associated Line Select Button. All actionsdescribed for memory locations 1 through 6 are also applicableto memory locations 7 through 12. If locations 1 through 6 arenot filled, the Second Memory Page will not be accessible.

− Color: Green.


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23 Code

01

RMU COM MEMORY PAGE


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24 Code


NAV MEMORY PAGE

1 - MEMORY PAGE IDENTIFICATION− Identifies the page as a NAV Memory Page.− Color: White.

2 - ACTIVE COM FREQUENCY− Identifies the NAV frequency that is currently active.− Color: White.

3 - NAV FMS STATUS ANNUNCIATION− In the NAV Memory Page, this field displays the FMS ENABLED

or DISABLED annunciation. This will be present whether or notthe Radio System interfaces with the FMS.

− Color: Yellow

NOTE: When the VOR or the ILS frequency is autotuned by theFMS, a magenta VOR or ILS frequency and a magentaAUTO label will be displayed on the top border of theRADIO Page NAV window.

4 - MEMORY PAGE SELECTED ANNUNCIATION− Indicates that the Memory Page is selected.− Color: Green.

5 - MEMORIES DISPLAY− Displays the preset frequencies and their associated locations.− When there is no frequency stored in a memory location only

the location number will be displayed in the associated memorydisplay line.

− Colors:− Memory identifications is green.− Frequency is cyan.

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25 Code

01

6 - MEMORY INSERT PROMPT− If it is desirable to insert a new frequency in a particular memory

location, simply press the Line Select Button beside the locationline, moving the yellow cursor box to that line. Then press theLine Select Button beside the Insert prompt label. This willcause all the data in memory from the insert location downwardto shift one position down. The cursor will remain in the insertionselected location allowing the new frequency to be tuned andstored in that memory location. A MEM FULL (Memory Full)annunciation will be displayed in the RMU display if the 12memory locations are filled and the Line Select Buttonassociated with the Insert prompt is pressed.

− Color: Green.

7 - MEMORY DELETE PROMPT− To delete a frequency from the memory, press the Line Select

Button adjacent to the line associated with the frequency to bedeleted. Then press the Line Select Button adjacent to theDelete prompt. The frequency enclosed by the cursor will bedeleted from the memory. Higher numbered memory locationswill then move upward to fill the empty memory location.

− Color: Green.



9 - MEMORY MORE PROMPT− The More prompt allows to display memory locations 7 through

12, by pressing the associated Line Select Button. All actionsdescribed for memory locations 1 through 6 are also applicableto memory locations 7 through 12. If locations 1 through 6 arenot filled, the Second Memory Page will not be accessible.

− Color: Green.


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26 Code



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27 Code

01

RMU NAV MEMORY PAGE


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2-18-11 Page

28 Code


ATC/TCAS CONTROL PAGE

1 - INTRUDER ALTITUDE DISPLAY− REL (green): Target’s altitude displayed relative to one’s own

airplane (default).− FL (cyan): Target’s altitude displayed as flight level (reverts to

REL after 20 sec).

2 - TA DISPLAY− AUTO (green): Traffic targets displayed only when TA or RA

target condition exists.− MANUAL (cyan): All traffic targets displayed within the viewing

airspace.

3 - FLIGHT LEVEL ID− Allows Mode S coding to reflect the current flight’s call sign. The

outer tuning knob moves the character position designator andthe inner tuning knob selects the desired alphanumericcharacter.

− Color: White

4 - FLIGHT LEVEL 1/2− Display of the transponder’s encoded altitude and the air data

source for that altitude.− Color: Green.



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29 Code

01

RMU ATC/TCAS CONTROL PAGE


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2-18-11 Page

30 Code


NAVIGATION BACKUP PAGE

NOTE: - The navigation information presented on the NavigationBackup Page are operationally identical to that normallypresented on the PFD.

- The compass card is presented only in arc partial format.

- The selected course and the DME distance to station areboxed.

- NAV and ADF active frequencies are also presented.

1 - ACTIVE NAV FREQUENCY

2 - BEARING 1 POINTER

3 - BEARING 2 POINTER

4 - ACTIVE ADF FREQUENCY

5 - COURSE DEVIATION BAR

6 - COURSE DEVIATION SCALE

7 - DME DISTANCE TO STATION

8 - MARKER BEACON DISPLAY

9 - SELECTED COURSE

10 - BEARING 2 SOURCE ANNUNCIATION

11 - BEARING 1 SOURCE ANNUNCIATION

12 - COMPASS CARD DISPLAY

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31 Code

01

RMU NAV BACKUP PAGE


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2-18-11 Page

32 Code


SYSTEM SELECT PAGE

1 - SYSTEM SELECT PAGE IDENTIFICATION− Identifies the SYS SELECT Page.− Color: White.

2 - COM 1 AND COM 2 BANDWIDTH SELECTION FIELD− Indicates the current COM 1 and COM 2 status regarding

bandwidth selection. Pressing the Line Select Button beside theCOM 1/COM 2 line field will toggle the receiver bandwidth fromWIDE (2 digits at the right of the decimal point) to NARROW (3digits at the right of the decimal point) or vice-versa.

− Color:− Cyan for COM 1 (2) BNDWD label.− Green for WIDE/NARROW indication.



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33 Code

01

RMU SYSTEM SELECT PAGE


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2-18-11 Page

34 Code


MAINTENANCE PAGE (POWER ON SELF-TEST)

1 - TEST PAGE IDENTIFICATION− Indicates where a failure has been detected.− Color: White.

2 - FAILURE SIDE IDENTIFICATION− Indicates the side of the detected failure.− Color: Green.

3 - FAILURE IDENTIFICATION− Identifies the detected failure according to the table below.− Color: Red.

ERRORMESSAGE MEANING ACTION DECISION

RMU ERR

One or moreinternalparameters weremeasured andfound to beoutside their self-test limit.

1. Check that CDH is not inEMERG Mode.

2. On main tuning page,perform tuning test on allradios by setting fre-quency and determiningthat radio is operating.

If tuning test fails,the RMU is notfully operable.

PRI BUS

Full RMU com-munications withall COMs, NAVs,and cross-sideRMU cannot beestablished on theprimary bus.

1. Check that all radiocircuit breakers are on.

2. Check RMU ON/OFFPage for all functionsON.

3. Check that CDH is not inEMERG Mode.

SEC BUS

Full RMU com-munications withthe on-side COMand NAV cannotbe establishedusing the secon-dary bus.

4. If 1 or 2 (or 3 if installed)are sources, correct andturn RMU power off for10 seconds. Reapplypower to start newPOST.

5. If error persists,

Any of thesemessages indi-cate that systemredundancy hasbeen reduced.

NAV UNIT/COM UNIT

The NAV unitsand/or COM unitscannot fullycommunicate withboth RMUs overprimary bus and/orthe on-side RMUover secondarybus.

perform on-side andcross-side tuning off allradios and activateauxiliary tuning sourcesto determine whichfunctions are stillavailable.

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35 Code

01

RMU MAINTENANCE PAGE (POWER ON SELF TEST)


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2-18-11 Page

36 Code


MAINTENANCE PAGE (PILOT ACTIVATED SELF TEST)

1 - SYSTEM TEST IDENTIFICATION− Indicates which unit is being tested.− Color: Amber.

2 - TEST RESULT INDICATION− Indicates whether the tested system is operating normally or not.− Color:

− Green for successful tests.− Red for unsuccessful tests.

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MANUAL



37 Code

01

RMU MAINTENANCE PAGE (PILOT ACTIVATED SELF TEST)


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2-18-11 Page

38 Code



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1 Code

01

TUNING BACKUP CONTROL HEAD

The Tuning Backup Control Head is a unit that provides an alternativemeans of tuning the NAV 2 and COM 2.

The TBCH is energized only when the AVIONICS MASTER is switchedON, and in normal operation it displays the RMU 2 NAV and COMactive frequencies (NAV 2 and COM 2).

NORMAL MODE

In the Normal Mode, the TBCH displays the RMU 2 NAV and COMactive frequencies. Each time these frequencies are tuned via RMU,the TBCH display is updated automatically. The same occurs whenthese frequencies are tuned via TBCH, the RMU 2 NAV and COMactive frequencies being also updated automatically.

It is also possible to tune the RMU 1 NAV and COM active frequenciesusing the RMU cross-side operational mode (see 2-18-11).

EMERGENCY MODE

When the TBCH is set to the Emergency Mode, the RadioManagement System will accept only the NAV and COM tuning viaTBCH, ignoring the RMUs control.

The RMUs will recover their capability of tuning the radio frequenciesonly when the TBCH is set to the Normal Mode again.

SELF TEST

After power up, the Tuning Backup Control Head performs a self-test.This test consists of saving the frequencies that the COM and NAVunits are tuned to as indicated by the Radio System Bus (RSB), andthen changing the frequency outputs to the COM and NAV andverifying that they have changed on the RSB. Failures are announcedin the display line associated with the function as an error messagefollowed by an error code “ERXX”, with the “XX” showing a two-digiterror code.This test is performed only on the ground, when the unit is turned on.


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2-18-13 Page

2 Code


TBCH CONTROLS AND INDICATORS

1 - SYSTEM INSTALLATION ANNUNCIATION− Indicates to which Radio System the Tuning Backup Control

Head is connected.

2 - TUNING CURSOR− Indicates which frequency may be changed by the Tuning

Knobs.

3 - NAV AUDIO ON ANNUNCIATION− Indicates that the NAV audio is selected on.

4 - EMERGENCY MODE ANNUNCIATION− Indicates when the unit has been selected to the Emergency

Mode, which inhibits RMU tuning capability.

NOTE: This annunciation is not related to the emergency COMfrequency of 121.5 MHz.

5 - SQUELCH ANNUNCIATION− Indicates that the squelch is opened by the SQ Switch.

6 - TRANSMIT ANNUNCIATION− Indicates that the COM transmitter is ON.

7 - NAV AUDIO BUTTON− Toggles NAV audio on and off.

8 - SQUELCH BUTTON− Toggles the COM squelch on and off.

9 - TUNING KNOBS− Change the frequency indicated by the tuning cursor.− Inner knob changes the frequency decimal digits in steps of

0.025 MHz for VHF and 0.050 MHz for VOR/LOC.

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3 Code

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It also changes the frequencies in the VHF sub-band thatcontains the 8.33 kHz spaced channels according to appropriateselection on the RMU. These frequencies are identified in voicecommunications by the channel names as exemplified below:

Frequency (MHz) Spacing Channel Name132,0000 25 132,000132,0000 8.33 132,005132,0083 8.33 132,010132,0166 8.33 132,015132,0250 25 132,025132,0250 8.33 132,030132,0333 8.33 132,035132,0416 8.33 132,040132,0500 25 132,050132,0... 8.... 132,...

− Outer knob changes the frequency non-decimal digits in steps of1 MHz for both VHF and VOR/LOC.

10 - NORMAL/EMERGENCY MODE SELECTOR BUTTON− When knob rotated clockwise selects normal Mode.− When knob rotated counterclockwise selects Emergency Mode.− EMRG button toggles the Emergency mode on and off.

11 - TRANSFER BUTTON− Alternately selects between the COM frequency (top) or the NAV

frequency (bottom) to be connected to the Tuning Knobs.− In the NAV only or COM only configurations, toggles the active

(top) frequency with the preset (bottom) frequency. In addition,holding the button down for two seconds will remove the presetfrequency and place the unit in the Direct Tuning Mode. Toreturn to the Active/Preset Tuning Mode, hold down the transferkey for two seconds.

12 - RADIO TUNING ANNUNCIATION− Identifies the frequency at the top and bottom lines.


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2-18-13 Page

4 Code



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5 Code

01

TUNING BACKUP CONTROL HEAD


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2-18-13 Page

6 Code


THIS PAGE IS LEFT BLANK INTENTIOANLLY

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1 Code

01

DIGITAL AUDIO PANEL

The EMB-135 BJ is equipped with three individual Digital Audio Panels(DAP), one each for the captain, copilot and observer.

This unit allows each flight crew member to select an individualtransceiver, the intercommunication function further permittingindividual selection and audio level adjustment of the followingcommunications equipment:

• VHF communication;• Crew/ramp station intercommunication;• Passenger address;• Reception and amplification of the NAV/COM audio signals.

NORMAL MODE

In the normal mode, each flight crew member may select one COMtransceiver (VHF COM 1, VHF COM 2, VHF COM 3 or HF), theinterphone function and, simultaneously, several audio receivers (COM1, 2 and 3, HF, NAV 1 and 2, ADF 1 and 2, and DME 1 and 2).

Also, the unit may provide volume control for each radio equipment,microphone selection between Boom and Mask (Oxygen Masks), andaudio output selection between Speakers and Headphones.

Other features are the capability to filter the NDB/VOR audio signals,attenuating morse code or voice signals. Finally, Normal Mode allowsmarker beacon audio sensitivity control, which also may silencetemporarily that type of signal.

EMERGENCY MODE

The emergency mode must be selected in case of Digital Audio Panelpower loss. In this case the captain will be directly connected to theCOM 1 and NAV 1 and the copilot to the COM 2 and NAV 2.

The interphone function will also be lost.

If power is recovered the Digital Audio Panel may be returned to thenormal mode of operation by selecting another MICROPHONE button(COM 1, 2, 3 or HF).


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2-18-15 Page

2 Code


COMMUNICATION SYSTEM SCHEMATIC

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3 Code

01

DIGITAL AUDIO PANEL CONTROLS AND INDICATORS

1 - MICROPHONE SELECTOR BUTTONS

− When pressed enables transmission and reception of radiosignals through the respective COM unit (COM 1, COM 2,COM 3, HF).

− Simultaneous selection of more than one microphone selectorbutton is not possible. Pressing a different microphone selectorbutton will cause the previously selected button to bedeselected.

− A bar illuminates inside the button to indicate that it is pressed.

2 - AUXILIAR BUTTON

− When depressed enables pilot´s communication through theSATCOM (if installed).

3 - PASSENGER BUTTON

− When pressed enables the crew to make the speech to thepassenger cabin while simultaneously deselecting the previouslyselected COM transmitter.

4 - EMERGENCY BUTTON

− In case of power loss to the Digital Audio Panel, connectsmicrophone directly to the emergency COM mic outputs andheadphone unit to COM and NAV audio.

− The captain is connected to COM 1 and NAV 1 and the copilotto COM 2 and NAV 2. Observer radio communications capabilityis lost.

5 - BOOM/MASK BUTTON

− Alternates selection between the boom (pressed) and the mask(released) microphones.

6 - ID/VOICE BUTTON

− When pressed (ID position), NDB and VOR audio signals arefiltered in order to enhance morse code identification.

− When depressed (VOICE position), VOR/ILS audio signals arefiltered in order to reduce morse code signal, enhancing theVOR/ILS voice associated messages (e.g., ATIS messages).


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01 REVISION 7

7 - HEADPHONE MASTER VOLUME CONTROL KNOB

− Allows adjustment of headphone amplifier volume.

8 - INTERPHONE SELECTOR KNOB

− When depressed, enables communications between captain,copilot, observer, and ramp station via airplane interphone.

− When rotated, provides interphone volume control.

NOTE: To enable the interphone function the respective controlwheel and observer communications switch must also beset at the HOT position.

9 - MARKER BEACON SENSITIVITY/MUTE KNOB

− The mute function is enabled by pressing the marker beaconsensitivity/mute knob and it is used to temporarily silence themarker beacon audio signal. The audio signal will beautomatically re-enabled according the following schedule:− If the mute function was selected when the marker beacon

audio level was above a certain threshold setting, the audiowill be re-enabled 5 seconds after the audio level descendsbelow that threshold setting.

− If the mute function was selected when the marker beaconaudio level was below that threshold setting, the audio signalwill be mute during 20 seconds, and then it will be re-enabled.

− The marker beacon sensitivity/mute knob, when rotated, alsocontrols the sensitivity of the Marker Beacon receiver.

10 - MARKER BEACON VOLUME KNOB

− When rotated, allows to control the marker beacon audiovolume.

NOTE: Does not allow volume settings below a certain level inorder to prevent the marker beacon audio from beingadjusted too low to be heard, that the marker signal couldbe missed.

11 - SIDETONE KNOB

− This knob selects the speaker ON (depressed) or OFF(pressed). It must be pressed when the headphones are used.

− The sidetone control is made by rotating the sidetone knob,which prevents undesirable feedback of speaker sidetoneaudio into the transmitting microphone.

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5 Code

01

12 - SPEAKER MASTER VOLUME CONTROL KNOB

− When rotated, allows adjustment of speaker volume.

13 - AUDIO CONTROL KNOBS

− When depressed, turns on the associated COM/NAV audio.− When rotated, provides volume control for the associated

COM/NAV audio.

NOTE: The COM 3 can be used for data transmission. When it isnot being used for voice transmissions, the COM 3 AudioControl Knob should remain pressed on all audio panels.


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2-18-15 Page

6 Code

01 REVISION 7

DIGITAL AUDIO PANEL

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1 Code

01

COMMUNICATION CONTROLS AND INDICATORS

COCKPIT

CONTROL WHEEL COMMUNICATIONS SWITCH (PTT)

1 - CONTROL WHEEL COMMUNICATIONS SWITCH

PTT POSITION - Momentary position. When pressed allows VHFand HF transmissions and speech to thepassengers through Passenger AddressSystem. Releasing this button, it returns to theHOT position and VHF, HF or passenger cabintransmissions will be interrupted.

NOTE: For VHF transmissions, a continuous command of PTTswitch is limited to 2 minutes. If the PTT switch is pressedlonger than 2 minutes, the message MIC STK will bedisplayed on RMU, and the microphone will be disabled.

HOT POSITION - Allows communication between crew membersand between crew members and ramp station.

OFF POSITION - Allows only audio reception.


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2-18-20 Page

2 Code

01 REVISION 6


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2A Code

01

CONTROL WHEEL


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2-18-20 Page

2B Code

01 REVISION 6

GLARESHIELD COMMUNICATION SWITCH (PTT)

1 - GLARESHIELD MIC PTT BUTTON− When pressed allows VHF and HF transmission and speech to

passengers through the Passenger Address System. Releasingthis button will interrupt transmission.

GLARESHIELD PANEL

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3 Code

01

CAPTAIN AND COPILOT HAND MICROPHONE

1 - HAND MIC PTT BUTTON− When pressed allows VHF and HF transmission and speech to

passengers through the Passenger Address System. Releasingthis button will interrupt transmission.

PILOT AND COPILOT CONSOLE


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4 Code


CAPTAIN AND COPILOT JACK PANELS

1 - CAPTAIN AND COPILOT JACKS− Allows to plug in headphone, boom microphone, and hand

microphone.

PILOT AND COPILOT JACK PANELS

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5 Code

01



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2-18-20 Page

6 Code


OBSERVER JACK PANEL AND COMMUNICATION SWITCH (PTT)

1 - BOOM JACK− Allows to plug in the boom microphone.

2 - HEADPHONE JACK− Allows to plug in the headphone.

3 - OBSERVER MICROPHONE SWITCHHOT POSITION - Allows communication with crew members and

ramp station.

OFF POSITION - Allows only audio reception.

PTT POSITION - Momentary position. When pressed allows VHFand HF transmissions and speech topassengers through the Passenger AddressSystem. Releasing this button, it returns to theOFF position and transmissions will beinterrupted, remaining only in audio reception.

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OBSERVER JACK PANEL


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RAMP STATION

FRONT AND REAR RAMP PANELS

1 - COCKPIT CALL BUTTON (momentary action)− When pressed, generates a tone in the headphones and cockpit

speakers.

2 - MICROPHONE/HEADPHONE JACK− Allows ramp crew to plug in a headphone and a microphone

equipped with a PTT Button.

NOTE: Ground crew panel is linked to the Hot Mic.

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01

FRONT AND REAR RAMP PANELS


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HF COMMUNICATION SYSTEM - KHF-950

The airplane may be equipped with a KHF-950 High-FrequencyCommunication System. All functions of the KHF-950 System arecontrolled by the KCU-951 Control Panel located at the controlpedestal.

HF OPERATING MODES

The KHF-950 High-Frequency Communications System provides thefollowing operating modes:

AMPLITUDE MODULATION

Amplitude modulation (AM) is a transmission process in which aselected frequency (called carrier frequency) and two sidebands(frequencies above and below the carrier) are generated andtransmitted. The upper sideband (USB) is the sum of the carrierfrequency and the voice, while the lower sideband (LSB) is thedifference between the two. The disadvantages of AM are that itoccupies a wide spectrum and is inefficient in the sense that a greatdeal of unneeded carrier is generated, as well as redundantinformation in the unused sideband.

SINGLE SIDEBAND

Single sideband operation achieves the same function as AM withconsiderably greater efficiency. The SSB transmitter electronicallyeliminates most or all of the carrier wave and one of the sidebands.The major advantages of SSB (either USB or LSB) as opposed to AMare greater talking power (about eight times that of AM for a givenpower input), reduced power drain, longer range and conservation ofthe spectrum (since only one sideband is required to transmit themessage).


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SUPPRESSED CARRIER AND REDUCED CARRIER (OPTIONAL)

The SSB operation with the carrier frequency eliminated is referred toas single sideband suppressed carrier and is designated as the A3Jmode in the KHF-950.If a small portion of the carrier frequency is transmitted along with thesideband, then the operation is referred to as single sideband reducedcarrier, and is designated as the A3A mode in the KHF-950.

SIMPLEX, SEMI-DUPLEX AND RECEIVE-ONLY OPERATION

Simplex operation means that the transmission and receptionfrequencies are the same. An example of simplex operation would becommunications with a control tower using a VHF COMM transceiver.Semi-duplex means transmit on one frequency and reception onanother frequency.Receive-only operation allows the system to operate as a receptoronly.The 99 user programmed channels can be programmed for eithersimplex, semi-duplex or receive only operation, and can operate in anyof the available modes (AM, USB or LSB).

NOTE: The use of LSB is legal for some international and off-shorecommunications, but is not authorized for use in the UnitedStates and most European countries.

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HF NORMAL OPERATION KCU-951 CONTROL UNIT

There are two types of operation:

- Discrete frequency tuning.

- Programmable channel.

DISCRETE FREQUENCY TUNING OPERATION

In the discrete frequency mode of operation, the user may directly tuneany one of 280,000 frequencies over the range of 2.0 to 29.9999 MHz.

1 - Access discrete frequency operation.

Apply power to the system by rotating the volume (V) knobclockwise from the OFF position. Wait for about two minutes, untilthe system has warmed up. Until then, no frequency is displayed.Make sure that the FREQ/CHAN button is depressed, in the FREQposition. Confirm this by seeing that no channel number isannunciated in the frequency display. If the button is pressed, amomentary press unlatches it and engages FREQ (direct tune)operation.

2 - Select the transmission mode.

Press the mode button to select the transmission mode (USB, LSBor AM).

3 - Enter the frequency.

Each digit in a frequency is selected individually. The largeconcentric knob on the lower right of the control unit may be rotatedin either direction and causes one of the displayed frequency digitsto flash. This flashing “cursor” indicates which frequency digit will bechanged by rotating the smaller concentric knob. Rotate the largerknob until the digit you wish to change flashes, and then select thedesired number into view by rotating the smaller knob.

4 - Tune the antenna.

Momentarily key the PTT to initiate the antenna coupler tuningcycle. During the tuning process the TX annunciator will flash andthe frequency numbers will blank. When the TX stops flashing andthe frequency reappears , the antenna tuning cycle is complete andyou are ready to transmit on the selected frequency.


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NOTE: - The discrete frequency mode always provides simplexoperation (transmit and receive frequencies are the same).

- Always key the PTT after selecting a new frequency to initiateantenna tunning, otherwise you may experience poorreception or miss an important call.

PROGRAMMABLE CHANNEL OPERATION

In the channel mode operation, the user may select user programmedchannels by their channel numbers. For user programmed channels:

1 - Access channelized operation.Apply power to system (rotate the VOLUME knob from the OFFposition). Wait for about two minutes, until the system has warmedup. Until then, no frequency is displayed. Make sure that theFREQ/CHAN button is pressed, in the CHAN position.

2 - Select the channel.Rotate the small inner concentric knob to select the desired channelnumber.

3 - Tune the antenna.Momentarily key the PTT to initiate antenna coupler tuning cycle.Adjust volume and squelch controls, as desired.

THE 99 USER CHANNELS PROGRAMMING PROCEDURE

The 99 user programmable channels available in the KHF-950 systemcan be programmed on the ground or in flight. All programmedinformation is stored in a nonvolatile memory and can be recalled byselecting the desired user channel number.

There are three types of channels that can be programmed:

1 - Semi-duplexThe user programs two different frequencies, one for receive andone for transmit. The user also assigns one of the availableoperating modes (USB, LSB or AM) to the selected channel. Semi-duplex operation is available only when the KHF-950 is beingoperated in the CHAN mode.

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2 - Simplex

The user programs the same frequency for receive and for transmit.The user also assigns one of the available operating modes (USB,LSB or AM) to the selected channel. Simplex operation is used byARINC, ATC (Air Traffic Control), and others.

3 - Receive-only

The user programs a frequency for reception and assigns one ofthe available operating modes (USB, LSB or AM), but leaves thetransmit portion of the preset channel blank.The transmitter and power amplifier are locked out and cannot beused when a channel has been programmed for receive-onlyoperation.

Receive-only channels are used to listen to frequency standards(W W V) for example, time, weather, Omega status, andgeophysical alert broadcasts.

SEMI-DUPLEX CHANNEL PROGRAMMING PROCEDURE

1 - Access channelized and program mode.

Apply power to the system by rotating the VOLUME (V) knobclockwise from the OFF position. Make sure that the FREQ/CHANbutton is pressed, in the CHAN position. Activate the program modeby pressing the PGM switch with a pointed object.

2 - Select the desired user channel.

Rotate the inner concentric knob to select the channel number to beprogrammed.

3 - Select emission mode.

Use the MODE button to select emission mode (USB, LSB or AM).Press the MODE button until the desired mode appears.

4 - Enter the receive frequency.

Use the outer larger concentric knob to position the flashing“cursor” on each digit of the receive frequency and use the smallerinner knob to select the desired number in each position.

5 - Store the receive frequency and mode of operation.

Push the STO button once and the receive frequency is entered inthe electronic memory. The TX annunciator will begin to flash in thedisplay window indicating the receive frequency is stored and youare ready to program the transmit frequency.


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6 - Enter and store the transmit frequency.

Use the outer larger concentric knob to position the flashing“cursor” on each digit of the transmit frequency and use the smallerinner knob to select the desired number in each position.

7 - Store the transmit frequency.

Push the STO button again and the transmit frequency is nowstored. If additional user channels are to be programmed, repeatsteps 2 through 7 at this time.

8 - Exit the program mode

Press the PGM switched with pointed object.


Momentarily key the PTT to initiate the antenna coupler tuningcycle. Adjust the volume (V) and squelch (S) controls, as desired.

SIMPLEX CHANNEL PROGRAMMING PROCEDURE

When you program a channel for simplex operation, both the receiveand the transmit frequencies will be the same. Programming a simplexchannel is similar to programming a semi-duplex channel, except theSTO button is pressed twice after the receive frequency and mode ofoperation are entered to store the frequency in both the receive andthe transmit positions.

RECEIVE-ONLY CHANNEL PROGRAMMING PROCEDURE

1 - Access channelized and program mode.

Apply power to the system by rotating the VOLUME (V) knobclockwise from the OFF position. Make sure that the FREQ/CHANbutton is pressed, in the CHAN position. Activate the program modeby pressing the PGM switch with a pointed object.

2 - Stow the cursor.

Stow the “cursor” if a frequency digit is flashing. The cursor isstowed by rotating the larger concentric knob until no frequencydigit is flashing. With the cursor stowed in the program mode, thesmaller inner knob is now used to select a channel number to beprogrammed.

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3 - Select channel number.

Use the smaller inner concentric knob to select the channel numberyou wish to program. Simply rotate the smaller knob until thedesired channel appears on the right side of the display.

4 - Select operating mode.

Push the MODE button until the appropriate mode (USB, LSB orAM) appears in the lower part of the display window.

5 - Set the frequency.

Move the “cursor” into position by rotating the larger concentricknob, and then use the smaller concentric knob to set the desirednumber under each “cursor” position.

6 - Store the information.

Push the STO (store) switch with a pointed object. The informationyou have entered will be stored in the electronic memory. After youpush the STO button, the TX annunciator will flash indicating thatthe frequency you have just entered is stored in the receiveposition, but nothing is entered in transmit. Since a receive-only isbeing set, ignore the flashing TX.

7 - Exit the program mode.

Press the PGM switched with pointed object. All information youhave stored is locked into CHANNEL memory.

FAULT INDICATION

If the system detects a fault during transmission or during the tuning ofthe antenna coupler, the frequency digits on the display begin to flash.Simply key the PTT button and the automatic antenna coupler begins anew tunning cycle to clear the fault.


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HF CONTROLS AND INDICATORS

KCU-951 CONTROL PANEL

1 - FREQUENCY DISPLAY

− Displays the operation frequency, in either FREQ or CHANmode.

2 - MODE DISPLAY

− Displays emission mode, transmit indicator and program modeindication.

3 - CHANNEL DISPLAY

− Displays the set channel. If in FREQ mode, the channel displaywill not show numbers.

4 - PHOTOCELL

− Automatically adjusts the display brightness.

5 - MODE SWITCH

− Selects transmission and reception mode. Momentarydepression cycles the KHF-950 from upper sideband (USB) tolower sideband (LSB) to AM.

6 - FREQ/CHAN SWITCH

− Allows the user to select between the two methods of frequencyselection. In the FREQ mode (depressed switch) the user maydirect tune any of the 280,000 available frequencies, simplexoperation only. In the CHAN mode (pressed switch), the userpresets the transmit and receive frequencies in up to 99available channels.

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7 - PROGRAM SWITCH

− Used on the 99 user-programmed channels programming.

8 - LARGER CONCENTRIC KNOB

− Moves the “cursor” which selects the digit of the frequency to bechanged.

9 - SMALLER CONCENTRIC KNOB

− Increments or decrements the frequency digit under the “cursor”(flashing digit).

10 - STORE SWITCH

− Used to store data when programming the user programmablechannels.

11 - POWER AND VOLUME KNOB

− Turns the system on and off and controls volume. Clockwiserotation past the first click turns the system on. Further rotationincreases audio level. After turning on, the system takes abouttwo minutes to warm up. Until then, no frequencies aredisplayed.

12 - SQUELCH KNOB

− Reduces background noises when rotated counterclockwise.Must be set by rotating the knob clockwise until backgroundnoise can be heard and then turning it counterclockwise untilbackground noise is eliminated or barely audible.

13 - CLARIFIER KNOB

− Clarifier is used only in SSB communications, and is notapplicable to AM mode;

− It must be used when due to off frequency ground stationtransmissions the audio voice quality from KHF-950 may soundunnatural;


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− To operate the clarifier, pull the knob out and rotate the knob ineither direction until the audio quality is optimized. When theknob is pushed in, the clarifier has no effect. When voice qualityis good and natural, the carifier knob should remain pushed in.

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01

KCU-951 CONTROL PANEL


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01

HF NORMAL OPERATION KFS-954 CONTROL UNIT

There are two types of operation:

- Discrete frequency tuning.

- Programmable channel.

DISCRETE FREQUENCY TUNING OPERATION

In the discrete frequency mode of operation, the user may directly tuneany one of 280,000 frequencies over the range of 2.0 to 29.9999 MHz.

1 - Access discrete frequency operation.

Apply power to the system by rotating the volume knob clockwisefrom the OFF position. Wait for about two minutes, until the systemhas warmed up. Until then, no frequency is displayed. Make surethat the EMMISSION MODE switch is in LSB, USB or AM position.The last frequency or the last channel used is displayed. In order tochange the last used channel into a discrete frequency, check firstto see that the display is showing a flashing zero in the channelposition or is blank in that position. If a channel number other thanzero is displayed, it will be necessary to move the “cursor” bydepressing the Frequency/Channel control knob. A digit on thedisplay will begin to flash. This flashing “cursor” indicates which digitin the display will be changed by twisting the Frequency/Channelcontrol knob. Each additional time you depress the knob the“cursor” will move one digit. Move the “cursor” until it is on thechannel number. Rotate the Frequency/Channel control knob untilthe channel number is set to zero.

2 - Select the transmission mode.

Select the EMMISION MODE switch to the transmission mode(USB, LSB or AM).

3 - Enter the frequency.

Press the Frequency/Channel control knob to cycle the “cursor”until the first digit in the frequency to be changed is flashing. (noticethat the channel number has changed from “0” to blank.) Rotatethis knob until have selected the desired number. Using theFrequency/Channel control knob in this same manner, change allthe digits necessary to display the desired frequency.


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01 REVISION 5

NOTE: The first one or two digits (MHz) of the frequency are displayedin the upper right of the display while the last four digits (kHz) ofthe frequency are displayed at the bottom of the display. Tostow the “cursor”, depress the Frequency/Channel control knobrepeatedly until no digit on the display is left flashing, or key thePTT momentarily.


Momentarily key the PTT to initiate the antenna coupler tuningcycle. During the tuning process the TX annunciator will flash andthe frequency numbers will blank. When the TX stops flashing andthe frequency reappears , the antenna tuning cycle is complete andyou are ready to transmit on the selected frequency.

NOTE: - The discrete frequency mode always provides simplexoperation (transmit and receive frequencies are the same).

- Always key the PTT after selecting a new frequency to initiateantenna tunning, otherwise you may experience poorreception or miss an important call.

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15 Code

01

PROGRAMMABLE CHANNEL OPERATION

In the channel mode operation, the user may select user programmedchannels by their channel numbers. For user programmed channels:

1 - Access channelized operation.

Apply power to the system by rotating the volume knob clockwisefrom the OFF position). Wait for about two minutes, until the systemhas warmed up. Until then, no frequency is displayed. Make surethat the mode selector knob is in the desired position (USB, LSB orAM) and that a channel number is showing in the display. If one isnot showing (discrete frequency tuning operation), or if you wish tochange the channel, move the “cursor” by pressing theFrequency/Channel control knob until the channel number isflashing.

2 - Select the channel.

Change the channel number by rotating the Frequency/Channelcontrol knob until the desired channel number appears. Thepreviously programmed receive frequency associated with thatchannel will appear in the display.


Momentarily key the PTT to initiate antenna coupler tuning cycle.Adjust volume and squelch controls, as desired.

THE 19 USER CHANNELS PROGRAMMING PROCEDURE

The 19 user programmable channels available with the KFS 954Control Display Unit can be easily programmed by the pilot on theground or in flight. Each of the 19 channels can be assigned aseparate frequency or frequencies (semi-duplex operation). Theoperating mode (USB, LSB or AM) of the stored channel is determinedby the position of the EMMISION MODE selector knob at the time thepilot is using the channel.


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01 REVISION 5

There are two types of channels that can be programmed:

1 - Simplex

The user programs the same frequency in receive and transmit.

2 - Semi-duplex

The pilot programs two different frequencies, one for receive andone for transmit.

All programmed information is stored in a nonvolatile memory and canbe recalled by selecting the desired user channel number.

SIMPLEX CHANNEL PROGRAMMING PROCEDURE

1 - Change the first frequency digit to enter the program mode.

Pushing the Frequency/Channel control knob move the “cursor” tothe first digit in the frequency to be changed. Twist theFrequency/Channel control knob until the desired number has beenselected for this frequency digit. As the knob is twisted, a flashingdash will appear to the right of the channel number to signify thatyou are in the program mode. You will be unable to receive ortransmit on the frequency displayed as long as the dash is flashingto indicate you are in the program mode. It is possible to change thelast digit (one tenth kHz) of the frequency without entering theprogram mode.

2 - Select the rest of the desired frequency.

Use the “cursor” by pressing the Frequency/Channel control knob toaddress each additional digit you want to change. Once the digit isflashing, again twist the knob to select the desired number.

NOTE: You may exit the program mode at any time and return to thepreviously stored frequency simply by keying the PPT.

3 - Store the frequency in the receive portion of memory.

Once the user have selected the desired frequency, press the STObutton to enter the displayed frequency in the receive portion ofmemory. The TX light will begin to flash indicating that memory isready to receive the transmit frequency.

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4 - Store the same frequency in the transmit portion of memory.

This is a simplex channel. Press the STO button a second time tostore the same frequency in the transmit portion of memory. Afterthe STO button is pressed the second time, the “cursor” will stowand the flashing dash will disappear to indicate the KFS 954 is nolonger in the program mode. Additional channels can beprogrammed by continuing to use the “cursor” andFrequency/Channel control knob in this manner.


Key the PTT to tune the antenna. After tuning is completed you areready to transmit.

SEMI-DUPLEX CHANNEL PROGRAMMING PROCEDURE

1 - Change the first frequency digit to enter the program mode.

Pushing the Frequency/Channel control knob, move the “cursor” tothe first digit in the frequency to be changed. Twist theFrequency/Channel control knob until the desired number has beenselected for this frequency digit. As the knob is twisted, a flashingdash will appear to the right of the channel number to signify thatyou are in the program mode. You will be unable to receive ortransmit on the frequency displayed as long as the dash is flashingto indicate you are in the program mode. It is possible to change thelast digit (one tenth kHz) of the frequency without entering theprogram mode.

2 - Select the rest of the desired frequency.

Use the “cursor” by pressing the Frequency/Channel control knob toaddress each additional digit you want to change. Once the digit isflashing, again twist the knob to select the desired number.

NOTE: You may exit the program mode at any time and return to thepreviously stored frequency simply by keying the PPT.

3 - Store the frequency in the receive portion of memory.

Once the user have selected the desired frequency, press the STObutton to enter the displayed frequency in the receive portion ofmemory. The TX light will begin to flash indicating that memory isready to receive the transmit frequency.

4 - Select the desired transmit frequency.


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01 REVISION 5

Use the “cursor” and the Frequency/Channel control knob again toselect a new transmit frequency.

5 - Store the transmit frequency in memory.

Press the STO button again to store the transmit frequency. Afterthe STO button is pressed the second time the “cursor” will stowand the TX light will go out. The flashing dash will also disappear toindicate the KFS 954 is no longer in the program mode. Additionalchannels may be programmed by continuing to use the “cursor” andFrequency/Channel control knob in the same manner.

6 - Press the PTT button to tune the antenna.

After tuning is complete you are ready to transmit. Before keyingthe PTT to talk, you may want to press the STO buttonmomentarily. This will allow you to listen momentarily to the transmitfrequency to avoid overriding someone else’s transmissions.

FAULT INDICATION

If the system detects a fault during transmission or during the tuning ofthe antenna coupler, the frequency digits on the display begin to flash.Simply key the PTT button and the automatic antenna coupler begins anew tunning cycle to clear the fault.

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HF CONTROLS AND INDICATORS

KFS-954 CONTROL PANEL

1 - OFF/VOLUME KNOB (inner concentric)

− Turns system on and adjusts audio volume.

2 - LAST FOUR DIGITS (kHz)

− Operating frequency are displayed in this area with EMISSIONMODE switch in LSB, USB or AM position.

3 - READOUT DISPLAY

− All frequencies and present channel numbers are displayed.

4 - USER PROGRAMED CHANNEL NUMBER DISPLAY

− Channel number is displayed in this area when using one of the19 programmable channels.

5 - DASH

− Indicates unit is in the PROGRAM MODE.

6 - FIRST ONE OR TWO DIGITS (MHz)

− With EMMISSION MODE switch in LSB, USB or AM position,the first one or two digits (MHz) of the operating frequency aredisplayed.

7 - TRANSMITION INDICATION

− Indicates unit is transmitting.

8 - PHOTOCELL

− Dims display automatically.

9 - STO (store) SWITCH

− Stores frequency in memory when pressed.


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10 - FREQUENCY/CHANNEL CONTROL KNOB (inner concentric)

− Allows the user to perform a variety of channel and frequencychanging functions. Depressing switch causes flashing "cursor"to move to the digit that the user desires to change. Appropriatefrequency or channel is then selected with rotary action. Thisswitch also serves as the clarifier function to adjust receivefrequency and improve speech quality in SSB operating mode.

11 - EMMISION MODE SWITCH

− Selects transmission and reception mode

12 - SQUELCH KNOB (outer concentric)

− Reduces background noises when rotated counterclockwise.Must be set by rotating the knob clockwise until backgroundnoise can be heard and then turning it counterclockwise untilbackground noise is eliminated or barely audible.

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01

KFS-954 CONTROL PANEL


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01 REVISION 5

HF SELECTION PANEL

Airplanes equipped with two HF equipment have a HF selection panelon the glareshield panel. This panel has a knob that allows to switchfrom HF1 and HF2 and vice-versa. This procedure turns off theprevious selected HF Control Panel turning automatically the other on.

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02

THIRD VHF COMMUNICATIONS SYSTEM -COLLINS 22AThe airplane may be equipped with a third VHF CommunicationsSystem. All functions of the Collins VHF-22A System are controlled bythe CTL-22 VHF Control Panel located at the control pedestal.

The Avionics Master DC Bus 1 supplies the third VHFCommunications System with a protective 5A circuit breaker.

THIRD VHF COM CONTROLS AND INDICATORS

1 - ACTIVE FREQUENCY DISPLAY− Displays the active frequency (frequency to which the equipment

is tuned) and diagnostics messages.

2 - XFR/MEM SWITCH− This is a 3-position, spring-loaded toggle switch.− When held to the XFR position, the preset frequency is

transferred up to the active display and the equipment retunes.The previously active frequency becomes the new presetfrequency and is displayed in the lower window.

− When held to the MEM position, one of the six stacked memoryfrequencies is loaded into the preset display.

− Successive pushes cycle the six memory frequencies throughthe display.

3 - FREQUENCY SELECT KNOBS− Two concentric knobs control the preset or active frequency

displays.− The large knob changes the digits to the left of the decimal point

in 1 MHz steps.− The smaller knob changes the digits to the right of the decimal

point in 0.005 MHz steps.− Numbers roll over at the upper and lower frequency limits.


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4 - ACTIVE BUTTON− Push the ACT button for about 2 seconds to enable the

frequency select knobs to directly retune the VHF-22A (activefrequency).

− The bottom window will display dashes and the upper windowwill continue to display the active frequency.

− Push the ACT button a second time to return the control to thenormal 2-display mode.

5 - TEST BUTTON− The self-test diagnostic routine is initiated in the transceiver by

pushing the TEST button.− The active and preset display intensity will flash, modulating its

brightness from minimum to maximum indicating self-test inprogress.

− The active frequency display will show four dashes and thepreset frequency display will show “00”.

− An audio tone will be heard from the audio system.− At the completion of the self-test program, the display will return

to its normal operation if no problem occurs.− In case of a detected failure, “diAG” (diagnostic) letters will be

displayed in the active and a 2-digit diagnostic code will bedisplayed in the preset display.

− Record any diagnostic codes displayed to help maintenancepersonnel in locating the problem.

6 - STORE BUTTON− The STO button allows up to six preset frequencies to be

selected and entered into the controls non-volatile memory.− After presetting the frequency to be stored, push the STO

button. The upper window displays the channel number ofavailable memory (CH1 through CH6); the lower windowcontinues to display the frequency to be stored. Forapproximately 5 seconds, the MEM switch may be used toadvance through channel numbers without changing the presetdisplay. Push the STO button a second time to commit thepreset frequency to memory in the selected location. Afterapproximately 5 seconds, the control will return to normaloperation.

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7 - POWER AND MODE KNOBOFF - Turns off the system.ON - Turns on the system.SQ OFF - Disables the receiver squelch circuits. Use this position

to set volume control or, if necessary, to try to receive avery weak signal that cannot operate the squelchcircuits.

8 - ANNUNCIATORS− The COM control contains MEM (memory) and TX (transmit)

annunciators.− The MEM annunciator illuminates whenever a preset frequency

is being displayed in the lower window.− The TX annunciator illuminates whenever the VHF-22A is

transmitting.

9 - PRESET FREQUENCY DISPLAY− Displays the preset (inactive) frequency and diagnostics

messages.− The frequencies displayed on the COM control show only five of

the six digits.

10 - COMPARE ANNUNCIATOR− ACT momentarily illuminates when active and preset

frequencies are being switched.− ACT flashes if the actual radio frequency is not identical to the

frequency shown in the active frequency display.


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5 Code

02

CTL-22 VHF CONTROL PANEL


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THIRD VHF COMMUNICATIONS SYSTEM -HONEYWELL RC-833/853The airplane may be equipped with a third VHF communicationssystem. All functions of the Honeywell RC-833/853 System arecontrolled by the RMU and by the DMU, if installed.

The active frequency will appear in the third COM window of the RMUdisplay when operating in voice mode. The desired frequency shouldbe selected by the pilot using the RMU select knob. Communicationcan be established pressing the third COM button on the audio paneland the PTT button.

When in data mode, the third VHF transmits the data on the frequencyselected by the DMU unit, while on the RMU’s third COM window theactive frequency will display “DATA”.

The third VHF has a pilot-activated self-test that performs a check ofindividual parts. A self-test is initiated when power is supplied to theRMU.

The Avionics Master DC Bus 1A supplies the third VHFCommunications System.

THIRD VHF COM NORMAL OPERATION

In order to operate the third COM unit for voice communications, it isnecessary to properly set the RMU. By pressing the PGE bezel buttonon the RMU, the page menu will be displayed. On the page menu the“SYS SELECT” option should be chosen. Once selected, the SystemSelect page will be displayed containing the available COM pairs (1/2,1/3 and 2/3), bandwidth of each COM unit and the COM3 operatingmode.

After selecting the COM pair to the corresponding RMU, the changebetween each COM mode of the selected pair is achieved by pressingthe “1/2” bezel button.

The switch between data and voice modes can be accomplished bymeans of the “SYS SELECT” page. The voice mode can also beentered selecting the COM3 window and turning the RMU knob. Thedata mode cannot be entered during a voice transmission. If a DMU isinstalled and connected to the third VHF, it will enter the data modeand “DATA” will be displayed in active frequency window when the thirdVHF starts a data transmission.


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It is possible to enter a frequency, store it, and perform a self-test usingthe same procedures applicable to regular COM units. Theannouncements also will be the same.

Should any radio component fail or a disagreement occur between thefrequency selection and the subsequent transmission, the frequency oroperating command will be removed from the RMU and replaced bydashes. This is an indication that the operation of that radio system isnot normal.

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RMU PAGE MENU


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SELCAL SYSTEMThe Ground-to-Air Selective Calling (SELCAL) System operates inconjunction with the communication radios. The SELCAL providescontinuous monitoring of a pre-set frequency, eliminating the need tocontinuously monitor the communication frequencies by the flight crew.

The SELCAL permits ground stations, equipped with encodingequipment, to call individual airplane by transmitting a coded signal.This coded signal will activate only one SELCAL unit to respond to thatparticular coded signal. In this case, a SELCAL voice message isactivated through the Aural Warning Unit. Once activated, the systemis reset for further monitoring by pressing the SELCAL Button, locatedon the Main Panel, or actuating the PTT function (on Control Wheel orglareshield panel).

NOTE: - For some airplanes the SELCAL enables only the VHF 2operation or only the HF operation.

- SELCAL will recognize the coded signal from ground stationsonly if the associated system (HF or VHF2) is powered onand its frequency is adjusted to the ground station frequency.


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SELCAL CONTROLS AND INDICATORS

1 - SELCAL BUTTON

− A striped bar illuminates inside the associated button to alert thecrew that communication is desired on VHF 2 or HF. A SELCALvoice message sounds simultaneously.

− When pressed, after a system activation, the striped barextinguishes and the system is reset.

SELCAL PANEL

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HONEYWELL MARK III COMMUNICATIONSMANAGEMENT UNIT (CMU)The airplane may optionally be equipped with Honeywell’s Mark IIICommunications Management Unit (CMU). Additional information onfunctions and operations can be obtained in the manufacturer’smanual.

The Mark III Communications Management Unit (CMU) is an airbornecommunications router that supports data link service access betweenthe following aircraft data link applications and their correspondingground service providers:

− Aeronautical Operational Communication (AOC)/Airborne FlightInformation System (AFIS).

− Controller Pilot Data Link Communication (CPDLC).− Automatic Dependent Surveillance (ADS).

The CMU is based on ARINC Characteristic 758 and can be upgraded,by software download, to an Aeronautical TelecommunicationsNetwork (ATN) router when protocols and application infrastructure areavailable to support Communications, Navigation and Surveillance/AirTraffic Management (CNS/ATM) data link applications. The CMUprovides an Aeronautical Radio, Inc. (ARINC) 724B compatible datalink router through which all character-oriented data are transmitted toand from the ground Aircraft Communications Addressing andReporting System (ACARS) network.

There are several levels of user interfacing. The CMU´s operationalcrew interface is provided through a Control Display Unit (CDU),printer, and data loaders.

Access to the ground network is provided via several ACARS air-ground sub-networks listed below:

− Satellite Communications (SATCOM).− Very High Frequency (VHF).− High Frequency (HF).− Mode-S (future).− Ultrahigh Frequency (UHF) (future).


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The CMU functions as a router, with growth capacity as a gateway andan end system for hosting data link applications. As a gateway, theCMU can implement future protocols and provide the appropriateprotocol conversions to support airborne systems whose architecturesdo not allow updates to future communication protocols but host thedata link applications.

Currently, the CMU is compliant with ARINC 758 level 0, which meansthat the CMU is able to perform ACARS functions. The other ARINC758 levels will be available in the equipment as soon as the softwareupgrades are released by the manufacturer.

The CMU is powered by the Avionic Switched DC Bus 1B.

The CMU interfaces with other airplane systems and equipmentthrough the following equipments:

− FMS The pilot → CMU interface isaccomplished through the Honeywell FMS(NZ-2000) Control and Display Units(CDU). It may be utilized with single ordual FMS installation.

− VHF #3 Transmit and receive data with groundbases.

− Printer Provides a hard copy data printout.

− Portable Data Loader Used to upload and to downloadinformation, customized configurationsand messages.

− CMC Sends maintenance information to theACARS.

− OOOI Used to inform if the airplane is in one ofthe following situations: OUT - parkingbrake released and doors closed;OFF - airplane lift-off (Weight Off Wheels);ON - airplane has landed (Weight OnWheels); IN - parking brake applied anddoors opened.

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CMU NORMAL OPERATION

Pilot interface is accomplished through the FMS Control andDisplay Unit (CDU). Data is entered in any field highlighted by acursor. Generally, when accessing a FMS page, the cursor will parkoff the page. Selecting the ENTER key will place the cursor overthe first enterable field on the page. Data is then entered into thefield with the alphanumeric keys.

In some situations, flight progress and related data from the FMSwill prefill into the field but it is possible to change that data byoverriding that value with a manual data entry.

When SEND is pressed, the message goes into a queue fortransmission. If the aircraft is flying over a region or is grounded atan airport not covered by a DSP (Data Link Service Provider)reception or transmission of messages to or from a ground stationwill not be successful.

Communications are eased by the use of the Main Menu page andseveral submenus to quickly access each function or serviceavailable.

CMU ABNORMAL OPERATION

The CMU has a Built-In Test (BIT) function that contains acontinuous monitor and self-test monitor. The continuous monitorfunction monitors critical system parameters and record faultsfound during normal operation.

The self-test monitor function is activated upon command from thesystem controller, external discrete input, front test panel switch,maintenance computer or a power-up event. The self-test monitorwill exercise various system functions, record faults found, andannunciate the results of the test.

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CMU CONTROLS AND INDICATORS

CONTROL DISPLAY UNIT (CDU)

1 - LINE SELECT BUTTONS− − There are four line selection buttons on each side of the Control

Display Unit (CDU) that provide selection of submodes withinmajor modes when in an indexed display.

2 - PREV/NEXT BUTTONS− PREV - Changes the current page to the previous page.− NEXT - Changes the current page to the next page.

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PRINTER CONTROLS AND INDICATORS

The CMU may be optionally equipped with a Printer. The Printer willprovide a hard copy printout of data collected by the CMU includinguplinked printer messages. For additional information on functions andoperations, refer to the manufacturer’s manual.

1 - PAPER ADVANCE− When pressed, advances paper.

2 - SELF-TEST− When pressed, it produces a test pattern, comprising full

character complement of printer, and activates both externalaural/visual alert functions. If button is pressed and held, printerwill successively print 40-, 60-, 80-column test patterns.Otherwise, each time the button is pressed, printer will print thenext test pattern in sequence.

3 - POWER ON− A green light indicator illuminates when power is on.

4 - ALERT RESET− Will reset the aural/visual alert function and causes Printer Busy

indicator to go out.

5 - PRINTER BUSY− An amber indicator illuminates upon receipt of first text

character and will remain on until the paper has advanced threelines beyond tear-off edge. The indicator will flash continuouslyafter paper advance until the reset button is depressed.

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COCKPIT VOICE RECORDERThe Solid State Cockpit Voice Recorder System records all audiosignals transmitted and received by the crew members via the DigitalAudio Panels, and any audible noise in the cockpit, through an areamicrophone installed below the standby compass.

The CVR is in operation whenever the essential DC Bus 2 isenergized, storing the last 2 hours of recorded information in a solidstate crash survivable memory unit. Any data older than 2 hours isautomatically overwritten by the most recent audio inputs.

A crash impact switch cuts off power to the CVR immediately afterexperiencing a 5 G impact in order to preserve the recorded data.

The CVR also incorporates an Underwater Locator Beacon (ULB).Powered by a dedicated battery, the ULB starts transmitting anacoustic signal in the 37.5 kHz frequency once it senses contact withwater, thus easing wreckage site location of a submerged airplane.The signal is transmitted during approximately 30 days.

A signal from the captain’s clock allows timing correlation betweenCVR and FDRS.

SELF TEST

When the TEST button is pressed the unit performs a functional self-test to verify the integrity of the system. A successful self-test results ina one-second activation of the status LED on the control panel and atwo-second 800 Hz tone that may be heard if a headphone is pluggedto the CVR control panel jack. If a failure is detected during the test,the status LED will not be activated and the 800 Hz aural tone will notbe heard.

ERASE FUNCTION

Previously recorded CVR data may be made unavailable if the ERASEbutton on the CVR control panel is pressed, provided the airplane is onthe ground and with the parking brake applied. In this case, only theCVR manufacturer will be able to recover the “erased” data.

When the ERASE button is pressed, a two-second 400 Hz tone maybe heard if a headphone is plugged to the CVR control panel jack,confirming that the erase command was successful.


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COCKPIT VOICE RECORDER CONTROLS ANDINDICATORS

1 - ERASE BUTTON− Erases previously recorded data from the crash survivable

memory.− Function is available only on the ground, with the parking brake

applied.

2 - TEST BUTTON− Tests system integrity.− A successful self-test results in a one second activation of the

status LED.− In case of failure, the status LED on the control panel is not

activated.

3 - HEADPHONE JACK− Allows plugging a headphone to monitor the 800 Hz test tone,

400 Hz erase tone and recorded audio signals.

4 - STATUS LED− Illuminates during one second to indicate a successful test.

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PASSENGER ADDRESS SYSTEMThe Passenger Address System (PAS) provides communication andannouncements from cockpit to the passenger cabin.

The PAS also interfaces with the audio entertainment and prerecordedannouncement systems to provide music and safety briefing/flightinformation through the passenger loudspeakers.

The following functions are available through the PAS:

− Voice announcement transmission (speech) to the PAX cabin.− Call function from passenger to attendant, through chime tone.− Chime tone for NO SMOKING and FASTEN SEAT BELTS signals.− Interface to boarding music and passenger briefing.

The PAS component responsible for sending signals to passengerentertainment and prerecorded announcement systems is thePassenger Address Amplifier (PAA), located in the airplane electroniccompartment.

The PAA establishes the priority among the input signals from theseveral sources and then drives these signals to the proper cabinloudspeakers. The PAA also provides the logic for generation of theaural and visual annunciators, chimes for passenger andNO SMOKING and FASTEN SEAT BELTS signals.


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PASSENGER ADDRESS OPERATING MODES

MUTED MODE

The Muted Mode is automatically selected during power up and whenno other mode is selected. In this mode there will be no chimes, nolights and no microphones enabled during power up or power supplytransients.

PILOT-TO-PASSENGER MODE

The Pilot-to-Passenger Mode is enabled by pressing the PassengerButton, labeled PAX, on the Digital Audio Panel. When this mode isenabled the captain, copilot or observer may transmit announcementsto the passengers, by pressing the respective PTT. The priority of thetransmission through the system is the following: captain, copilot,observer. There are no chimes in this mode.

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SATCOM SYSTEM

INTRODUCTION

The airplane may be equipped with a TT-3000 Series Aero-MSATCOM, that is capable of performing global communication throughthe INMARSAT satellite network. The system provides one channelvoice 4800 BPS AMBE coding, fax 2400 BPS on 2 wire RJ-11 typeinterface, circuit mode data 2400 BPS, and can be interfaced with aPC via a RS-232 serial port for data communication.

NOTE: The information presented in this manual has the purpose ofassisting the user to perform basic operations on the SATCOMsystem. For advanced operation refer to the manufacturer’suser manual.

SATCOM OPERATION

When the Aero-M system is powered-up, after a short pause thehandset will display the initialization page.

Once that has ocurred, the display will shift to the Wait for GPS page.This means that the system is acquiring GPS (Global PositioningSystem) satellite signals and determining the systems position.

The display will then shift to the Wait for NCS (Network CoordinatingStation) page, indicating that the system is attempting to logon to asatellite network and acquire a bulletin board.

Finally, the display will then shift to the Logon Display and the user isnow ready to proceed with Pre-Operational Requirements.


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PRE-OPERATIONAL REQUIREMENTS

Before attempting to initiate or receive a call, the user must verify thatthe Aero-M System has logged on to the satellite network. There aretwo indications that the system has power applied and the system haslogged onto the satellite network:

− Both the left and green handset LED indicators will beilluminated and the signal strength indicator will display ameasurement;

− The Aero-M logon display will appear in the handset display.

The top line of the display will indicate which Ocean Region Satellite isin use followed by a colon. This will be:

− AORW : Atlantic Ocean Region West Satellite;− AORE : Atlantic Ocean Region East Satellite;− POR : Pacific Ocean Region Satellite;− IOR : Indian Ocean Region Satellite.

NOTE: The Aero-M system requires the user to determine the oceanregion in which the aircraft is currently located. The user mustenter the user menu and setup the system for that oceanregion.

SELECTION OF OCEAN REGION SATELLITE

From Logon display, or Wait for NCS display, enter Area Menu, bypressing “2nd” and “3” keys. The “∗” indicates the current oceanregion selected. Scroll through menu to the correct region, using the“↓” key. Select correct ocean region pressing the “OK” key.

The display will revert to Wait to NCS page and when NCS is acquiredthe display will return to Logon Display with the new ocean regionused.

After the colon, the display will indicate what Land Earth Station (LES)is in use for routing your calls. If the LCD display is on another display,you can always return to the Logon display by pressing the “Exit” keyrepeatedly to step back from MENU operation. Many functions requirestarting from the Logon display.

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OPERATION

This section contains information on types of calls and procedures fordialing and completing calls.

GROUND CALLS

All ground-to-air calls must be placed as international calls. EachINMARSAT Satellite has an assigned three digit code that is, for allpractical purposes, the country code for the satellite. It is necessary tofirst dial the international access code that is used in your country,followed by the satellite country code and the number issued by theGround Service Provider for the Airborne Earth Station (AES). In somecountries, call routing by the Public Switched Telephone Network canbe routed to the “870” country code. “870” is the Single NetworkAccess Code (SNAC). SNAC is a special country code that contactsthe Network Coordinating Station for INMARSAT Aero-M services andautomatically routes the call to the satellite that Airborne Earth Stationis logged on to.

In countries where calls do not route to 870, it is necessary to manuallyinclude the appropriate country code for the INMARSAT Satellite that isserving the Airborne Earth Station. If the incorrect satellite/countrycode is dialed, an automatic response will inform that the station youare calling is not currently logged onto this ocean region. If you aresure the AES is logged on to a satellite, you can redial the numberusing another satellite country code in order to contact the AES.

Country codes for the INMARSAT Satellites:

AREA/OCEAN REGION INTERNATIONALACCESS CODE

Single Network Access Code SNAC 870

Atlantic Ocean Region East AORE 871

Pacific Ocean Region POR 872

Indian Ocean Region IOR 873

Atlantic Ocean Region West AORW 874


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VOICE CALLS

− Receiving calls

If a handset is receiving a call, the handset will ring and theYellow Ring LED will flash. To establish a connection the usercan either press the Toggle Hook key (key 5 on HANDSET KEYFUNCTIONS figure) or lift handset from cradle. When aconnection has been established the Yellow Ring Led will beilluminated.

− Standard telephone calls

Dial a telephone number by entering call prefixes, country andarea codes and number, then press the “#” key or the ToggleHook key (key # 5 on HANDSET KEY FUNCTIONS figure) inthe handset keypad. To finish the call press the the Toggle Hookkey (key # 5 on HANDSET KEY FUNCTIONS figure).

− Calls from handset #2 and #3

Handset #2 can initiate short code, service code, handset tohandset, as well as standard telephone calls.

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SATCOM CONTROLS AND INDICATORS

The TT-3000 Series Aero-M SATCOM System uses the TT-5621AHandset as the main user interface. The handset contains a keypadwith 21 separate function keys, a 2X12 character Liquid Crystal Display(LCD), four indicator LEDs, and a volume control.

The handset allows the user to:− Navigate the user menu;− Place and receive calls;− Save and recall phonebook entries;− Configure system parameters;− View phone log;− Monitor system operations;− View system errors as they occur.

HANDSET #1 KEYS FUNCTIONS

The handset #1 keypad consists of 21 keys (12 numeric and ninefunction keys). These keys are designed to perform multiple functions,that are dependent on the selected mode.

The handset has three modes that determine the function of the keys:− Normal mode: keys perform their primary functions;− Alpha mode: keys perform alphanumeric functions;− 2nd mode: keys perform their alternate functions.


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The following table lists all the handset #1 keys and defines theirfunctions:

MODEFigureNumber NORMAL ALPHA 2nd

1 Accept selection

2 Toggles handsetaudio mode

3 Enable alphamode

Disable alphamode Delete entry

4 Move oneselection up Edit entry

5 Toggle hook

6 Move oneselection down Mute

7 38 29 610 511 912 813 #14 015 ∗16 717 4

18 1- ? ! , . : “ ‘ $ ( ) + /1

Access usermenu

19 Enable 2nd mode Enable 2nd mode Disable 2ndmode

20

Delete/Backspace onecharacter /Delete display(hold 2 seconds)

Insert entry

21Exit menu /Cancel selection

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HANDSET #2 AND #3 KEYS FUNCTIONS (OPTIONAL)

The handset #2 and #3 are optional and are installed on the connectedin parallel. Their keypad consists of 21 keys (12 numeric and ninefunction keys). These keys are designed to perform multiple functions,that are dependent on the selected mode.

The handsets #2 and #3 have two modes that determine the functionof the keys:

− Normal mode: keys perform their primary functions;− Alpha mode: keys perform alphanumeric functions.

The following table lists all the handset #2 and #3 keys and definestheir functions:

MODEFigureNumber NORMAL ALPHA

1 Transfer call2 Memory3 Memory location 34 Memory location 25 Toggle hook6 Microphone mute7 3 D E F 38 2 A B C 29 6 M N O 610 5 J K L 511 9 W X Y Z 912 8 T U V 813 #14 015 ∗16 7 P Q R S 717 4 G H I 418 1 - ? ! , . : “ ‘ $ ( ) + / 119 Last number redial20 Memory location 121 Storage

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HANDSET #1 LCD DISPLAY

The handset uses a LCD to display all data. The LCD is a 2X12character alphanumeric display. Additionally, the LCD will display tensymbols to indicate operating status and alert the user of any errorsthat may occur. The following table lists all display symbols, referring tothe figure in the next page, and defines their meaning.

FIGURENUMBER

NAME DESCRIPTION

1 More entries above flag

Indicates that additionalentries are availableabove and can bedisplayed by pressingthe “Edit” key.

2 New AES messages flag

A recorded MESmessage for theSDU/SIM is waiting atthe LES.

3 Signal strength indicatorIndicates the strength ofthe signal beingreceived by the system.

4 Hook off flagIndicates that a call is inprogress.

5 Speaker active flagThe speaker in thecradle is active.

6 Toggle field flag

Indicates that pressing“↑” or “↓” keys canchange the currentselection.

7 Alpha mode flag

Indicates that the “Del”key was pressed andthe alpha mode isactive. The next keypressed will perform itsalpha mode function.

Continued

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FIGURENUMBER

NAME DESCRIPTION

8 Security enabled flag

Indicates that thehandset has beenlocked and can only beaccessed by entering avalid PIN code.

9 2nd mode flag

Indicates the “2nd” keywas pressed and the2nd mode is active. Thenext key pressed willperform its alternatefunction.

10 More entries below flag

Indicates that additionalentries are availablebelow and can bedisplayed by pressingthe “↓” key .


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The figure below shows all the symbols found on the handset display.

HANDSET DISPLAY


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HANDSET #1 LED INDICATORS

The handset uses four colored LEDs as indicators. These LEDs alertthe user of the following indications:

FIGURENUMBER

LED DESCRIPTION

1 Left Green Indicates power tohandset.

2 Right Green

System is logged or islogging into the satellitenetwork. When on, theantenna strengthindicator will show ameasurement.

3 Yellow

When on, indicates callin progress, the LED willflash when a call isreceived.

4 RedIlluminates when anerror has occurred in thesystem.

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HANDSET #1 LED INDICATORS


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HANDSET #2 AND #3 LED INDICATORS

The handset uses two colored LEDs as indicators. These LEDs alertthe user of the following indications:

FIGURENUMBER

LED DESCRIPTION

1 Left GreenIndicates power tohandset.

2 Right Green Ring indicator.

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HANDSET #2 AND #3 LED INDICATORS


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POTS TELEPHONE ADAPTER (OPTIONAL)

The pots telephone adapter provides the interface between theSATCOM system and the airplane cockpit headset audio system. Thisequipment provides remote dialing capability for the SATCOM system.

This interface allows the cockpit to maintain full headset utilization forSATCOM calls or normal headset operation at the touch of the AUXbutton on the pilot’s or copilot’s digital audio panel.

The pots telephone adapter keypad consists of 17 keys (12 numericand 5 function keys).

The following table lists all pots keys and their functions.

FigureNumber CONTROL FUNCTION

1 HOOK Hook switch key2 13 24 35 46 57 68 REDIAL Redial key9 VOL Volume control key10 FLASH Flash key11 # Send key12 *13 014 915 816 717 HOLD Hold key

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IRIDIUM SATELLITE TELECOMMUNICATIONSYSTEMThe airplane may be equipped with an AirCell mobile transmitter,ST 3100 Iridium-based telecommunication system. The ST 3100telephone accesses the Iridium Satellite Network composed of 66 LowEarth Orbit (LEO) satellites and operates in the frequency range of1616 MHz to 1625.5 MHz. When in use, the telephone system controlsthe power level at which the phone transmits. The transmission powerlevel of the ST 3100 can range up to a maximum of 7 watts.


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01 REVISION 4


HANDSET

1 - OPERATING DISPLAY AND STATUS INDICATORS

− Provide a visual indication of telephone numbers and systemoperational status.

2 - InUse

− Illuminates when the phone is in use.

3 - NoSvc

− Illuminates when the telephone is not registered.

4 - VOLUME KEYS

− Adjust ear piece volume

5 - VOLUME/SCROLL DOWN KEY

− Scrolls Down/Volume Down..

6 - VOLUME/SCROLL UP KEY

− Scrolls Up/Volume Up.

7 - DOT IN LOWER LEFT CORNER OF DISPLAY

− Illuminates when the phone is on.

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01 REVISION 4

OPERATION

POWER ON

Anytime the airplane is energized the phone has power. During initialpower up, a small dot ‘’•” appears in the lower left corner of the display.The dot will remain as long as the phone has power. The wordIRIDIUM will also appear in the display. During this time, the wordsNoSvc (No Service) will briefly appear at the bottom center of thedisplay. However, a call cannot be placed until the NoSvc light isextinguished.

NOTE: Power off can only be achieved by turning off the main avionicsswitch.

PLACING A CALL FROM THE AIRPLANE

1 - With the power on and the NoSvc light out, enter the number thatyou would like to call (USA/Canada Only-001+Area Code+Number)OR (International Calls-00+Country Code+Number).

2 - Press the SND key.3 - When a call has been placed an InUse message will illuminate at

the bottom left of the display. The InUse message will remain untilthe phone call has been terminated by pressing the END key.

4 - The caller may hear up to five “beeps” in the 10-12 secondsrequired for the system to complete the call.

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RECEIVING A CALL

When the telephone rings, remove the handset from the cradle andpress the SND key to answer the phone.

END A CALL

Always press the END key to terminate the current call.

AUTOMATIC REDIAL

To redial the last number that was called, press the SND key and thenumber will be redialed but will not appear on the display.

PLACING A CALL TO THE AIRPLANE FROM THE GROUND

NOTE: - Placing a call to an AirCell ST 3100-equipped airplanethrough the Iridium Network is simple and easy. All calls toand from Iridium-based telephone systems are consideredinternational calls; in fact, the Iridium Satellite Network has itsown Country Code. Therefore, when calling an Iridium-equipped airplane, always use 8816 as the Iridium CountryCode.

- Direct-dialed calls to an AirCell ST 3100-equipped airplaneare billed by long distance providers as an international call.These rates can vary greatly. Callers may wish to consulttheir long distance provider for rate information prior tomaking a call. As a cost-effective alternative to Direct-Dialing,callers may wish to use Two-Stage Dialing that is explainedbelow.

- Some long distance providers do not recognize the IridiumSatellite Country Code of 8816. In addition to providing theleast-cost method of calling an AirCell ST 3100-equippedairplane, Two-Stage Dialing will recognize the Iridium CountryCode.


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01 REVISION 4

- In the following procedures, **00 represents the InternationalDirect Dial (IDD) access code prefix of the country from whichthe call is being placed. In many countries, the InternationalDirect Dial (IDD) code is 00. However, this is not always thecase. For the most current and an extensive listing of codes,please consult the Iridium Web site page:www.iridium.com/customer and refer to the User Manualsection or one of the many Internet sites that contain thisinformation.

DIRECT DIALING

− Placing a call from a ground phone (USA/Canada) to an Iridium-equipped airplane, dial: 011 (International Direct Dial access codeprefix) + 8816 (Iridium Country Code) + XXX.XXXXX (the X’s denotethe Iridium Network Number). Caller may hear a pause of up to 25seconds as the system completes the call.

− Placing a call from a ground phone (outside the USA/Canada) to anIridium-equipped airplane, dial: **00 (International Direct Dial accesscode prefix of the country from which the call is being placed) + 8816(Iridium Country Code) + XXX.XXXXX (the X’s denote the IridiumNetwork Number). Caller may hear a pause of up to 25 seconds asthe system completes the call.

TWO-STAGE DIALING

Two-Stage Dialing may provide the most cost-effective method ofcalling an AirCell ST 3100-equipped airplane. Utilizing this dialingmethod, the calling party is charged by their long distance telephoneprovider for a long distance call to Tempe, Arizona, USA for theduration of the call. The called party (in the airplane), is charged AirCellper minute charges for airtime from the time the call connects inTempe, Arizona, USA until the call is terminated.

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− When placing a call from a ground phone to an AirCell ST 3100-equipped airplane, dial the automated Iridium Call RoutingCenter 480.768.2500. Note that the Call Routing Center islocated in Tempe, Arizona, USA. As a result, long distancecharges will be billed to the caller as if they were dialingsomeone in Tempe, Arizona, USA.− Placing a call from a ground phone (USA/Canada) to the

Iridium Call Routing Center, dial: 1.480.768.2500.− Placing a call from a ground phone (outside the

USA/Canada) to the Iridium Call Routing Center, dial: **00(International Direct Dial access code prefix) + 1 (USACountry Code) + 480.768.2500.

− An automated recording will answer and instruct the caller to dialthe Iridium Phone Number 8816 + XXX.XXXXX.

− The call will be routed to the AirCell ST 3100-equipped airplane.If the system is registered on the network, in four to six secondsthe system will ring. The called party may receive the call bypressing the SND key on the handset.

− After completing the call, the user may end the call by pressingthe END key.

PLACING A CALL TO ANOTHER IRIDIUM PHONE

To place an Iridium-to-Iridium call, dial 00 + 8816 + XXX.XXXXX (othersubscriber’s Iridium phone number).


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INERTIAL REFERENCE SYSTEM (IRS)The Inertial Reference System (IRS) is responsible for generatingattitude and heading data on EMB-135 BJ. The airplane is equippedwith two identical and independent IRS. The dual IRS installation isused as an additional navigation sensor for the FMS, and it may alsoprovide accurate attitude and heading data to the Head-Up GuidanceSystem (HGS), when this equipment is installed.

Another possibility for HGS-equipped airplanes is the installation of asingle IRS tailored for that equipment, while PFDs, MFDs and RMUsreceive information from two AHRS.

Basically, the IRS provides attitude and heading reference data for theairplane displays and navigation data for the FMS by sensing linearmotion and angular rates through inertial sensors.

The IRS interface with the airplane’s systems and equipment is asfollows:

− Air Data Computers (ADC 1 and ADC 2): The IRS 1 and IRS 2receive altitude, altitude rate and true airspeed information from theADC 1 and ADC 2 respectively, to improve the precision of computednavigation data.

− Integrated Computers (IC1 and IC2): The IRS 1 and IRS 2 providepitch, roll and heading information to the respective PFD, andheading information to the respective MFD, through the IC-600s.Data is transmitted separately to both sides, to ensure that a singleIC failure does not compromise the data path.

− Radio Management units (RMU 1 and RMU 2): the IRS 1 providesheading information to both RMUs via DAU 2.

− Autopilot System: The IRS 1 provides pitch, roll and accelerationinformation to the Autopilot System via IC-600-1.

− Weather Radar: The IRS 2 provides attitude information to theWeather Radar for antenna stabilization.

− Flight Management System (FMS): The IRS provides attitude,heading and navigation information to the FMS. The IRS receivesposition, magnetic heading, UTC time and date from the FMS.

− EGPWS: The IRS 1 provides attitude and heading information to theEGPWS.


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− Stall Protection System (SPS): The IRS provides attitude ratevariation and vertical acceleration information to the SPS.

− Integrated Standby Instrument System (ISIS): The IRS 1 providesheading information to the ISIS.

− Head-Up Guidance System HGS: The IRS 1 (or the dedicated IRS)provides acceleration, speed, attitude, heading and wind informationto the HGS.

− Windshear Detection And Escape Guidance System: The IRS 1provides attitude rate variation and vertical acceleration informationto the windshear computer.

− Flight Data Recorder (FDR): The IRS 1 provides attitude andheading information to the FDR via DAU 2 and IC-600.

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IRS INTERFACES WITH OTHER SYSTEMS


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INERTIAL REFERENCE SYSTEM COMPONENTS

Each IRS consists of an Inertial Reference Unit (IRU) and a ModeSelect Unit (MSU).

Each IRU uses two 28 VDC power inputs, one for normal power(primary source) and the other for backup power (airplane batteries).The IRU 1 primary power source is the Essential DC Bus 1 and itsbackup power source is the Backup Essential Bus. The IRU 2 primarypower source is the DC Bus 2 and its backup power source is theBackup Bus 2. If the IRU loses the primary power, it automaticallyswitches to backup power.

When the IRU operates solely on backup power, it will operate for 30minutes and the ON BATT annunciator on the Mode Selector Unit willilluminate. The IRS 1 (2) ON BATT advisory message will bepresented on the EICAS.

INERTIAL REFERENCE UNIT (IRU)

The Inertial Reference Unit contains three laser gyros and threeaccelerometers that are mounted on each of the three axis inside theIRU, which it uses to measure inertial motion.

The IRU requires initialization data from the Flight ManagementSystem and Air Data Computer. From the inertial measurements,initialization data, and air data inputs, the IRU performs the calculationsnecessary to provide position, velocity, heading and attitude data to theairplane.

The IRUs are located in the forward electronics compartment.

MODE SELECTOR UNIT (MSU)

The Mode Selector Unit is a cockpit-mounted control panel thatprovides mode selection, status indication and test initialization for theassociated IRU. It has a four-position mode select switch and statusand fault annunciators.

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IRS OPERATING MODESThe IRS operating modes may be selected by setting the MSU modeselect switch to the desired mode. Under certain conditions the IRUmay automatically revert to specific modes and sub-modes.

ALIGNMENT MODES

MSU SET FROM OFF TO ALIGN OR NAV

In the alignment mode, the IRU aligns its reference axis to the localvertical and true north, and estimates the horizontal earth ratecomponents to compute latitude. The latitude at which the IRU isaligned affects the alignment time. The relationship between alignmenttime and latitude is shown in the chart below.

0

5

10

15

20

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80

AL

IGN

ME

NT

TIM

E -

min

ute

s...

..

ALIGNMENT LATITUDE - degrees Northern and Southern

The airplane must remain stationary during alignment, while the MSUALIGN annunciator is lit and IRS 1 (2) ALN advisory messagepresented on the EICAS. If the IRU detects excessive airplane motionIRS 1 (2) EXC MOTION advisory message is presented on the EICASand it starts an automatic full realignment 30 ± 1 seconds after themotion stops. Normal passenger-loading or cargo-loading activitiesshould not cause an excessive airplane motion condition.


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NOTE: To complete the alignment, the IRU requires a valid input ofthe airplane’s present position (latitude and longitude) from theFMS.

While the present position is not entered on FMS the IRS 1 (2) NOPPOS caution message will be presented on EICAS. If the presentposition is not entered during the normal alignment time, the MSUALIGN annunciator flashes and the IRS 1 (2) ALN FAULT cautionmessage will be presented on the EICAS. The IRU will not completethe alignment phase and/or will not enter the NAV mode until itreceives a valid position input from the FMS.

The IRU accepts multiple entries of latitude and longitude. More thanone entry may be necessary to confirm, update or correct the position.However, the IRU does not accept new position inputs until 2 secondsafter the previous input. A new position entry overwrites the previousentry, and the IRU uses only the latest entry for its calculations. TheIRU conducts a position comparison test on latitude and longitudeimmediately after each data has been entered. To pass the test, theentered data must compare within 1 degree of the storedlatitude/longitude from the last power down from the NAV mode. If thetest fails, the MSU ALIGN annunciator flashes and an IRS 1 (2) ALNFAULT caution message will be presented on the EICAS. If a secondpresent position entry do not pass in the test the IRS 1 (2) FAIL will bepresented on the EICAS.

No attitude and heading is displayed during align mode.

MSU SET FROM NAV TO ALIGN OR NAV TO ALIGN AND BACK TONAV (TO BE USED ON THROUGH FLIGHTS):

This command will revert the IRU to the align down-mode, whichzeroes residual velocity errors accumulated during the previousnavigation mode operation. The fine leveling process of the align modeis reactivated, and heading accuracy is updated. When the IRU entersthe align down-mode, the ALIGN annunciator illuminates and an IRS 1(2) ALN advisory message is presented on the EICAS.

The airplane must be stationary (ground speed less than 20 knots)during the align down-mode otherwise an automatic full realignmentwill begin in 30 seconds after an excessive aircraft motion stops.

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AUTOMATIC NAVIGATION REALIGNMENT (ANR):

The ANR is entered automatically and concurrently with normal modeoperations, when the IRU senses that the airplane has completed a“no motion period” (between 7.5 and 15 minutes, as a function oflatitude). In the ANR sub-mode, the IRU applies corrections by zeroingvelocities, resetting the attitude platform, and correcting heading errorsaccumulated during the previous navigation mode operation. When theIRU enters the ANR sub-mode, the ALIGN annunciator illuminates andan IRS 1 (2) ALN advisory message is presented on the EICAS.

NAVIGATION MODE

The IRU enters the NAV mode after completing its alignment when theIRU MSU switch is set to NAV. In the NAV mode, the IRU uses the lastvalid position data entered during the align mode or align down-modeas its initial present position and updates the present position basedonly on inertial data while it remains in the NAV mode. The IRUalgebraically adds computed magnetic variations from a magneticvariation topographical map (MAGVAR) to true heading and true trackto produce magnetic heading and track magnetic angle. The magneticheading and magnetic tracking angle outputs are set to no computeddata (NCD) inside a northern and southern latitude cutout area.

ATTITUDE MODE

The attitude mode is the IRU’s reversionary mode. It is automaticallyentered by the IRU if it experiences in-flight loss of power, or it may beselected by the crew if the FAULT annunciator lights and an IRS 1 (2)FAIL caution message is presented on the EICAS, indicating a criticalfault, that invalidates all outputs.

CAUTION: THE MSU SWITCH MUST NOT BE SET TO ATTITUDEMODE INADVERTENTLY. IF THIS OCCURS ON THEGROUND, IT IS NECESSARY TO SET THE MSUSWITCH TO OFF FOR AT LEAST 3 SECONDS, THENBACK TO ALIGN OR NAV. IF ATTITUDE MODE IS SETIN-FLIGHT, IT IS NOT POSSIBLE TO RECOVER THENAV MODE.


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This mode provides a quick attitude restart: during the first 20 secondsthe IRU enters the erect attitude transitional mode. In this transitionalmode the IRU computes a new level axis set. The airplane must beheld steady, straight and level until the MSU ALIGN annunciatorextinguishes itself and an IRS 1 (2) ALN advisory message disappearsfrom the EICAS.

When operating in the attitude mode the IRS 1 (2) ATT MODE cautionmessage is presented on the EICAS. In this mode, attitude outputs arenot as accurate as when operating in the NAV mode, and magneticheading must be entered periodically through the FMS Control DisplayUnit.

POWER-OFF MODE

The IRU enters the power-down mode when the MSU switch is set toOFF and remains in that position for 3 seconds (to prevent inadvertentselection). After that, power continues for 10 to 15 seconds to transferthe last calculated latitude and longitude and other IRS parameters toits non-volatile memory. The ALIGN annunciator illuminates when theIRU is in the power-down mode (and an IRS 1 (2) ALN advisorymessage is presented on the EICAS).

NOTE: The airplane must not be de-energized before the ALIGNannunciator light extinguishes.

De-energizing the airplane before the ALIGN annunciator extinguishesitself (and an IRS 1 (2) ALN advisory message is removed from theEICAS) may interrupt the transferring process of the IRU’s lastcalculated position and other IRS parameters to its non-volatilememory, which will affect the next alignment.

TEST MODE

The test mode is selected by pressing the MSU test switch. The testmode can be selected in either the align mode or the NAV modewithout affecting basic IRS function. The test mode is inhibited in theattitude mode and in the NAV mode when aircraft ground speedexceeds 20 knots.

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MAGNETIC VARIATION LATITUDE CUTOUTS


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IRS OPERATING PROCEDURES

POWER-ON AND ALIGNMENT

− Check the MSU switch set to OFF.− After energizing the airplane press the AVIONICS MASTER Buttons.− Set the MSU switch to ALIGN or NAV. ALIGN annunciator

illuminates. (ON BATT and NO AIR annunciators may illuminatebriefly). An IRS 1 (2) ALN advisory message is displayed on theEICAS.

− Through the FMS initial position page, accept one of the presentposition options.

− Wait for completion of alignment. When selecting NAV directly, theIRU automatically enters the NAV mode after successful completionof alignment.

NAVIGATION MODE ENTRY

− If the MSU switch was set to NAV, make sure that the ALIGNannunciator goes out and an IRS 1 (2) ALN advisory messagedisappears from the EICAS.

− If the MSU switch is set to ALIGN, set the MSU switch to NAV whenthe NAV RDY annunciator lights.

− The IRU outputs inertial data to all displays and systems.

ALIGNMENT DOWNMODE (“FAST ALIGNMENT” - to be used onthrough flights)

− The airplane must be stationary.− Pull the MSU switch out of NAV detent and set it from NAV to

ALIGN. The ALIGN annunciator will illuminate and an IRS 1 (2) ALNadvisory message displayed on the EICAS.

− If the IRU position is the same as the known position, set MSUswitch to NAV after the ALIGN annunciator illuminates and an IRS 1(2) ALN advisory message is presented on the EICAS. The IRUautomatically enters NAV mode after 30 seconds.

− If the IRU position is different from the known position, re-initializethe IRU with a new position entry through the FMS CDU.

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ATTITUDE MODE SELECTION

The following procedure applies to the selection of the ATT mode afterthe FAULT annunciator has illuminated and an IRS 1 (2) FAIL cautionmessage is presented on the EICAS:− Set the MSU switch to ATT.− If the FAULT annunciator extinguishes itself, the ALIGN annunciator

will illuminate for 20 seconds and an IRS 1 (2) ALN advisorymessage is presented on the EICAS.

− Hold the airplane steady, straight and level until the ALIGNannunciator extinguishes itself and an IRS 1 (2) ALN advisorymessage disappears from the EICAS.

− Initialize the IRU with magnetic heading through the FMS CDU.− Update heading periodically through the FMS CDU.− When operating in the attitude mode the IRS 1 (2) ATT MODE

caution is presented on the EICAS.

CAUTION: WHEN OPERATING IN THE ATTITUDE MODE,ATTITUDE OUTPUTS ARE NOT AS ACCURATE ASWHEN OPERATING IN THE NAV MODE, ANDMAGNETIC HEADING MUST BE ENTEREDPERIODICALLY, THROUGH THE FMS CDU.

POWER DOWN

− Set the MSU switch to OFF. The ALIGN annunciator illuminates andan IRS 1 (2) ALN advisory message is presented on the EICAS.

− After the ALIGN annunciator extinguishes itself and an IRS 1 (2) ALNadvisory message is removed from the EICAS, de-energize theairplane.


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IRS EICAS MESSAGES


IRS 1 (2) OVERHEAT The associated IRS isoverheated.

IRS 1 (2) ATT MODE The associated IRS isselected to attitude mode.

CAUTION

IRS 1 (2) ALN FAULT The associated IRS did notcomplete the alignmentphase successfully.

IRS 1 (2) FAIL The associated IRS hasfailed.

IRS 1 (2) NO PPOS The present position hasnot been entered.

IRS 1 (2) NO MAG HDG There is no magneticheading available.

IRS 1 (2) ALN The associated IRS is inthe alignment phase or theIRU mode select switch isset at ALIGN position.

ADVISORY IRS 1 (2) ON BATT The associated IRS isbeing powered by theairplane batteries.

IRS 1 (2) EXC MOTION The associated IRSdetected excessive motionduring the alignment phase.

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IRS CONTROLS AND INDICATORS

MSU CONTROL PANEL

1 - TEST SWITCHTEST − − When the TEST switch is pressed the IRU enters

the self-test sub-mode.

2 - ANNUNCIATORSALIGN − − Indicates that the IRU is in the align mode. A

flashing ALIGN annunciator indicates thatlatitude/longitude was not accepted by the IRU.

FAULT − − Indicates an IRU fault.ON BATT − − Indicates that the IRU is being powered by the

airplane batteries.BATT FAIL − − Indicates that the airplane batteries are incapable

of sustaining IRS operation on backup poweroperation.

NAV RDY − − Indicates that alignment is complete, if the MSUswitch is set to ALIGN.

NO AIR − − Indicates that cooling air is inadequate to cool theIRU.

3 - MODE SELECTOR SWITCH− OFF: The IRU circuitry is OFF. However, when the IRU is

operating and the MSU switch is moved from another position toOFF, there will be a delay of 3 seconds before the IRU starts thepower-down routine, in order to prevent its inadvertent selection.

− ALIGN: Once this mode is selected the IRU starts the fullalignment or alignment down-mode. ALIGN annunciator is ONand an IRS 1 (2) ALN advisory message is presented on theEICAS.

− NAV: There is a detent in the NAV position. The switch must bepulled out of the detent to be moved to another position.

− ATT: If the MSU switch is placed at this position, after 2 seconds(time delay incorporated to prevent inadvertent selection of theATT mode) the IRU enters the erect attitude sub-mode for 20seconds, during which the ALIGN annunciator illuminates. Thenthe IRU enters the attitude mode and the IRS 1 (2) ATT MODEcaution message will be presented on the EICAS. On the ground,the MSU switch must be set to OFF for at least 3 seconds beforethe ALIGN or NAV mode can be re-established.

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IRS MODE SELECTOR UNIT


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IRS INDICATIONS ON THE PFD

ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI)

1 - ATTITUDE SPHERE− Color:

− Sky: blue.− Ground: brown.

2 - ROLL SCALE− Color: White− Range: 360 degrees.− Resolution: 10, 20, 30 and 60 degrees for left and right roll

attitudes.− Fixed pointers (unfilled triangles) are located at zero degrees

and 45 degrees (LH and RH).

3 - ROLL POINTER− Color: White.− Provides the roll angular indication against the roll scale.

4 - EXCESSIVE PITCH CHEVRONS− Color: Red− Marks –45 and 65 degrees pitch up, and 35, 50 and 65 degrees

pitch down.

5 - PITCH SCALE− Color: White.− Range: 0 to 90 degrees (pitch up and pitch down).− Marks:

− Pitch up: 0, 5, 10, 15, 20, 25, 30, 40, 60 and 90 degrees.− Pitch down: 5, 10, 15, 20, 25, 30, 45, 60 and 90 degrees.

6 - GROUND/SKY REFERENCE EYEBROW− Color: Blue or brown.− The eyebrow provides a quick ground/sky reference for attitudes

where the horizon line is out of the display.

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ELECTRONIC ATTITUDE DIRECTOR INDICATOR


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ATTITUDE DECLUTTER

When there is an excessive attitude situation, certain indicators areremoved in order to declutter the PFD.

Excessive attitude situation occurs when roll attitude is greater than 65degrees, or pitch attitude greater than 30 degrees nose up or20 degrees nose down.

In this case, the following symbology shall be removed from thedisplay:

− Flight Director couple arrow,− Low Bank limit arc,− Flight Director command bars,− Vertical Deviation scale, pointer and label,− Radio Altitude digits, label and box,− Marker beacons indicators,− Decision Height digits and labels,− Selected Airspeed bug and indicators,− Vertical Speed bug and indicators,− Selected Altitude bug, indicators and box,− All failure flags associated with the items listed above,− The Heading, Radio Altitude, LOC, GS, and ILS comparison monitor

displays.

The PFD indicators will be restored when the two conditions below aremet:

− Roll attitude less than 63 degrees left and right.− Pitch less than 28 degrees nose up and greater than 18 degrees

nose down.

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ELECTRONIC HORIZONTAL SITUATION INDICATOR (EHSI)

1 - COMPASS CARD DISPLAYMay be displayed in the Full Compass or Arc formats, selected viathe Display Control Panel (see section 2-18-40).− Color: white.− Range: 360 degrees.− Resolution: 5 degrees.

2 - HEADING LUBBER LINE (FULL COMPASS FORMAT)− Color: White.− Provides the current heading reading against the heading scale.

3 - CURRENT HEADING DIGITAL DISPLAY (ARC FORMAT)− Color:

− Open box: white− Digits: white

− Range: 0 to 360 degrees.− Resolution: 1 degree.

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EHSI - FULL COMPASS AND ARC FORMATS


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COMPARISON MONITORS

1 - ATTITUDE COMPARISON MONITOR DISPLAY− Label: ROL, PIT or ATT.− Color: Amber.− If roll information deviates by more than 6 degrees between the

PFD 1 and PFD 2, a ROL comparison monitor will be displayedinside the attitude sphere.

− If pitch information deviates by more than 5 degrees betweenthe PFD 1 and PFD 2, a PIT comparison monitor will bedisplayed in the upper-left portion of the attitude sphere.Simultaneous activation of the both pitch and roll comparisonmonitors will be announced by an ATT label displayed in theupper-left portion of the attitude sphere, in the same field of theROL and PIT comparison monitors.

2 - ATTITUDE FAILURE DISPLAY− Removal of the pitch scale and roll pointer.− Coloring the attitude sphere overall blue.− A red ATT FAIL label is displayed on the top center of the

attitude sphere.

3 - ATTITUDE SOURCE ANNUNCIATION− Label: ATT1 for IRS 1 and ATT2 for IRS 2.− Color: Amber when one AHRS supplies both sides or both IRS

are supplying cross-side.− Annunciations are removed when both IRS are supplying on-

side PFDs.

4 - HEADING SOURCE ANNUNCIATION− Label:

− MAG1 or MAG2 when cross-side IRS source is selected.− Color:

– For MAG: amber when the same IRS is supplying both sidesor both IRS are supplying cross-side.

− When both IRS are supplying on-side, annunciation is removed.− If a heading source becomes invalid the heading source

annunciation will refer to the invalid heading source, HDG1 orHDG2, as applicable.

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5 - HEADING COMPARISON MONITOR DISPLAY− Color: Amber.− Label: HDG− Activated when a difference of 6 degrees between both PFDs is

found and the airplane roll is less than 6 degrees.− For airplane rolls greater than 6 degrees, annunciation will be

displayed if the difference between both PFDs is greater than12 degrees.

− The HDG threshold will be restored to 6 degrees if the airplaneroll is less than 5 degrees for 90 seconds. Otherwise, a 12degrees HDG threshold will be maintained.

6 - HEADING FAILURE DISPLAY− Digital heading bug symbol is removed and a red HDG FAIL

annunciation is displayed on the PFD and MFD compass cards.− The bearing pointers, map display, To/From, selected heading

bug, drift angle, selected course/track and course deviationdisplays will be removed.

− Heading source annunciation will be HDG 1 or HDG 2.− Heading select and course select/desired track digital display

will be replaced by amber dashes.

7 - COURSE DEVIATION FAILURE− Pointer is removed.− Red X displayed over the scale.


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IRS FAIL INDICATION ON THE PFD


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FLIGHT MANAGEMENT SYSTEM (FMZ 2000)The FMZ 2000 Flight Management System (FMS) controls a completerange of navigation functions. Its primary purpose is to provide highaccuracy in long range lateral and vertical navigation. The system maybe installed with a single or dual configuration. Should the airplanehave a dual configuration, each unit can provide navigation data to theother unit. For additional information on functions and operation, referto the manufacturer’s manual.

The FMS is mainly composed of the following components:

− Control Display Unit (CDU).− Navigation Computer (NZ).− Data Loader (DL) or Portable Data Transfer Unit (PDTU).

The FMS operates in the following situations: Oceanic, Remote, NorthAtlantic Minimum Navigation Performance Specification Airspace,Enroute, Terminal, Non-Precision Approach and Required NavigationPerformance 10.

The FMS interfaces with the followings systems and equipment:

− GPS sensor(s), ADC 1 and 2 - The GPS receives satellite datathrough the passive GPS antenna, processing and blending collecteddata with ADC data and sends the resulting information to the FMScomputer.

− IRS 1 and 2 - Provides the necessary data to compute wind and forDead Reckoning Mode, when the subsystem is not capable ofnavigating by itself.

− MFD and PFD - The FMS provides data for display navigationguidance on the PFD and navigation map data on the MFD.

− RMU 1 and 2 - The RMU interfaces with the FMS computer tocontrol the operating frequencies, modes and channels of thevarious radios. For the dual configuration, each RMU supplies eachrespective on-side NZ.

− COM 1 and 2, NAV 1 and 2 - The FMS includes a radio-tuning pageon which the pilot can manually select the VHF NAV and COMfrequencies. Only the NAV frequency is fed back to the FMScomputer for verification of the tuning action. COM 1 and 2 interfacewith FMS through the RMUs. The FMS can also automatically selectthe NAV radio frequencies.

− The FMS also provides latitude and longitude to TCAS.

The Control Display Unit (CDU), located on the control pedestal,provides control functions management and operating modes forproper FMS operation.


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FMS SCHEMATIC


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FMS OPERATING MODES

FMS FUNCTIONS

NAVIGATION

The navigation function computes the airplane position and velocity forall phases of flight. The navigation priority modes, based on sensoraccuracy, are as follows:

− GPS− DME/DME− VOR/DME− IRS (if installed)

The GPS is the most accurate sensor. When the GPS is in use, theother sensors are still monitored for position differences, but they donot contribute to FMS position, unless the GPS becomes inaccurate,unavailable or is manually deselected. In this case, the FMSautomatically tunes the DME/DME in order to provide position. WhenDME/DME is not accurate, the VOR/DME is selected.

On airplanes equipped with dual Inertial Reference System (IRS),replacing the AHRS, the IRS is used as a primary navigation sensorwhen other navaid are not available.

If all position sensors and radios are lost, the FMS shifts to DegradeMode (DGRAD) and in approximately 2 minutes it enters the DeadReckoning Mode (DR). In this mode, the position is calculated usingthe last known airplane position. The ground speed and track areestimated with IRS heading, ADC TAS and the last known wind data.

The dual FMS configuration may operate with dual IRS and dual GPSproviding four long-range navigation sensors. The sensors status maybe accessed in the NAV INDEX 1/2 page.

In this configuration, on-side FMS outputs and flight plan informationare available to the opposite-side FMS through an interconnecting bus.

The automatic tuning is made through the RMU for computing anoptimum position. The FMS also includes a radio-tuning page on whichthe pilot can manually select VHF NAV, COM, ADF and transponderfrequencies. The FMS has the capability of tuning communicationfrequencies in the 8.33 kHz channel spacing.

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FLIGHT PLANNING

The flight planning function computes the active flight plan with bothlateral and vertical definition.

When the FMS long-range navigation is selected, the flight directorcommand bars will provide the visual command to bank the airplane tothe desired track.

The VNAV is applicable only for the descent path and it is not coupledto the flight director, being only a reference information displayed onthe PFD glide slope scale.

Additionally the navigation computer can be programmed by theoperator to automatically fly different types of holding patterns.

DATA BASE

The database contains worldwide coverage of navaids, airways,departure procedures, approach procedures, Standard TerminalArrival Routes (STARs), airports and runways. This information isupdated every 28 days. The database can also store up to 200 pilot-defined flight plans and waypoints, which are only updated whenchanged by the pilot.

In single configuration, the Data Loader (DL) is used to update theDatabase, transferring data to and from the Navigation Computer. Inthis configuration, this unit can be installed on the left lateral console,close to the pilot’s mask stowage box.

In dual configuration, the Portable Data Transfer Unit (PDTU) is usedto reload entire information package at each update by using a 3 1/2"floppy disk.

NAVIGATION DISPLAY

A multiple waypoints map, based on the airplane’s present position,can be displayed on the MFD. It comprises the Waypoints connectedby white lines defining a pre-planned route, and also navaids andairports.


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FMS MODES

The dual FMS configuration provides four operating modes that maybe accessed through the FMS MAINTENANCE 1/3 page:

DUAL MODE

In this mode, the following information is automatically transferred tothe cross-side FMS: flight plan, performance data, waypoints definedby the pilot, flight plans created in one system and radio tuning.

NOTE: For the proper operation in DUAL mode it is necessary to usethe same software version, same NAV and CUSTOM databases and same settings for both systems in the ConfigurationModules. The initial position difference between both systemsshall not be more than 10 NM.

INITIATED TRANSFER MODE

In this mode the flight plan and performance data entry will only betransferred to the cross-side FMS through the prompt commandavailable in the last page of the ACTIVE FLT PLAN pages. Waypointsdefined by the pilot, created flight plans and radio tuning areautomatically transferred to the cross-side FMS.

NOTE: For the proper operation in INITIATED TRANSFER mode it isnecessary to use the same software version, same NAV andCUSTOM data bases and same settings for both systems inthe Configuration Modules. The initial position differencebetween both systems shall not be more than 10 NM.

INDEPENDENT MODE

In this mode, only the radio tuning is automatically transferred to thecross-side FMS.

NOTE: To operate in the INDEPENDENT mode, it is necessary to usethe same software version and same settings in theConfiguration Modules. If any of these requirements is notaccomplished, the system automatically passes for thepossible operating mode. For instance, if only the CUSTOMdatabase differs in both systems, the operating modeautomatically switches from DUAL to INDEPENDENT.

SINGLE MODE

No information is exchanged between both systems.

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FMS CONTROLS AND INDICATORS

CONTROL DISPLAY UNIT (CDU)

1 - ANNUNCIATORS

The annunciators are located on the top of the LCD display.− Colors:

− White: indicating advisory annunciation.− Amber: indicating alerting annunciation.

DSPLY (White) Illuminates when the CDU displays a page thatis not relative to the current airplane lateral orvertical flight path. This annunciator is notshown on the PFD.

DR (Amber) Illuminates when a radio updating loss occurs,as well as all other position sensors, for a periodlonger than 2 minutes.

DGRAD (Amber) Illuminates when the FMS cannot guarantee theposition accuracy for the present phase of theflight.

MSG (White) Illuminates when there is a message (advisoryor alert) on the scratchpad. The annunciatorturns off when the message is cleared from thescratchpad.

OFFSET (White) Illuminates when a lateral offset path has beenentered in the FMS. The annunciator turns offwhen the offset is removed.

APRCH (White) Illuminates when the FMS is selected asnavigation source and the following conditionsare valid: a non-precision instrument approachhas been activated from the navigationdatabase, the airplane position is between 2 NMoutside the final approach fix and the missedapproach point, the DGRAD must be off andFMS using approved sensors for non-precisionapproach.

NOTE: The FMS transmits all the annunciators to the PFD, except theDSPLY annunciator, so the pilot must not trust only on the FMSCDU for checking the FMS system status.


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2 - LINE SELECT BUTTONS

− There are four line select buttons on each side of the CDU thatprovide the following functions:− Select submodes within major modes when in an indexed

display.− Used as direct access to the other FMS modes when in a

non-indexed display.− Enter data to the scratchpad.

3 - BRIGHTNESS CONTROL BUTTON

− Used to manually control the brightness of the display.− Using this button, the photo sensors are activated and maintain

the brightness level through a wide range of lighting conditions.The brightness is adjusted pressing up or down the Bright/Dimbutton, and a control bar will be displayed in the scratchpad.

− The brightness can be adjusted so that, during daylightconditions, the display cannot be seen.

4 - MODE BUTTONS

PERF Displays the performance pages.

NAV Displays the NAV index pages.

FPL It may be used to display the first page of the activeflight plan, if the flight plan was previously entered, tomanually create a flight plan, to select a stored flightplan and to create a flight plan for storage.

PROG Displays the first progress page, the current status ofthe flight.

DIR Displays the active flight plan with the DIRECT andINTERCEPT prompts.

5 - ALPHANUMERIC BUTTONS

− Consist of alphabet letters, the numbers 0 through 9, a decimaland a slash. It is used to enter inputs to the FMS. ASP (Space) key is used to insert a space following a characterin the scratchpad, and a +/- (Plus/Minus) key will result ina - being entered, changing to + in a subsequent press.

− The alphanumeric keys make entries only on the scratchpad.

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6 - FUNCTIONS BUTTONS

PREV Changes the current page to the previous page.

NEXT Changes the current page to the next page.

CLR Clears alphanumeric entries in the scratchpad or ascratchpad message.

DEL Works together with line select buttons in order to deletewaypoints and other items displayed on the CDU. Thisbutton is inhibited when a message is displayed.

The CDU has five function buttons directly above the LCD display thatwill not work if pressed. The following messages will be displayed inthe scratchpad:

VIDEO VIDEO NOT AVAILABLE.GRAPHIC GRAPHIC NOT AVAILABLE.ATC ATC NOT AVAILABLE.BACK BACK COMPLETE.FN FN NOT AVAILABLE.

7 - SCRATCHPAD

− It is the working area, located on the bottom line of the display,where the pilot can enter data and/or verify data before lineselecting the data into its proper position.

− Data is retained on the scratchpad throughout all mode andpage changes.

− The scratchpad also provides advisory and alerting messages tobe displayed.

The colors are designed to highlight important information. Colorassignments are coordinated as much as possible with other displays.See below the parameters associated to each color:

Vertical Cyan (Blue)Atmospheric Data Cyan (Blue)Lateral GreenFROM Waypoint YellowTO Waypoint MagentaPrompts and Titles WhiteFlight Plan Names OrangeIndex Selections Green


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FMZ 2000 FMS CD-820 CONTROL PANEL


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JOYSTICK (OPTIONAL)

The joystick functions are available through the joystick controller thatis located on the control pedestal and through the selection of the MFDJSTK menu.

When the MFD joystick menu is selected, the joystick controller isavailable to control the Designator Symbol movement on the MFDFMS flight plan.

JOYSTICK OPERATION

On power-up, the designator is co-located with the present flight planwaypoint position.

If MAP mode is selected, moving the joystick controller, will cause theDesignator Symbol to be displayed in blue color with a broken linewhich moves in the same direction from its last waypoint position.

If PLAN mode is selected, moving the joystick controller, the flight planmoves to the opposite direction from its last position, while theDesignator Symbol remains fixed at the center of the plan format.

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JOYSTICK CONTROLLER


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JOYSTICK MENU BUTTONS FUNCTION AT MAP MODE

− SKIP ("SKP") button: Skips the designator to the position of the nextwaypoint in the flight plan in case of the designator is co-locatedwith a plan waypoint. Otherwise, the designator broken line tailskips to the next waypoint in the flight plan.

− RECALL ("RCL") button: Positions the designator at the presentposition of the airplane and removes the designator box from thedisplay in case of the designator is co-located with the flight planwaypoint. Otherwise, the designator is positioned over the waypointfrom which the designator line is extended and the designator line isremoved from the display.

− ENTER ("ENT") button: The latitude and longitude coordinates ofthe designator are transmitted to the selected FMS scratchpad as arequested waypoint.

JOYSTICK MENU BUTTONS FUNCTION AT PLAN MODE

− SKIP ("SKP") button: Positions the flight plan so the next waypointis displayed over the designator in case of the designator is co-located with a flight plan waypoint. Otherwise, skips the tail of thedesignator line to the next waypoint in the flight plan.

− RECALL ("RCL") button: Positions the designator at the presentposition of the airplane and removes the designator box from thedisplay in case of the designator is co-located with a flight planwaypoint. Otherwise, it positions the designator over the waypointfrom which the designator line is extended and removes thedesignator line from the display.

− ENTER ("ENT") button: The latitude and longitude coordinates ofthe designator are transmitted to the selected FMS scratchpad as arequested waypoint.

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MFD JOYSTICK MENU


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NAVIGATION DISPLAYSThe navigation data provided by the Radio Management System andFlight Management System are displayed to the crew through thePFDs, MFDs and RMUs.

ADF and/or VHF NAV bearings and VHF NAV or FMS CDI (CourseDeviation Indicator) are displayed on the PFD in an ElectronicHorizontal Situation Indicator (EHSI). The EHSI navigation sources aswell as the display format (Full Compass or Arc) may be selected bythe crew via the Display Control Panel (DCP).

Several other navigation data are also presented on the PFDs: GS(Glide Slope) pointer, DME distance, Ground Speed/Time-to-go,marker beacon indicators, wind intensity and direction vector, etc.

The MFDs present Weather Radar, TCAS and the route selected onthe FMS. Additional information is also presented on the MFD: windintensity and direction vector, TAS, Time-to-go, etc.

The RMUs NAV Backup Page also present the EHSI, in the Arc formatonly (see section 2-18-11).


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DISPLAYS CONTROLS AND INDICATORS

DISPLAY CONTROL PANEL (DCP)

1 - DISPLAY FORMATS SELECTOR BUTTON− Pressing the FULL/WX Button alternates the EHSI presentation

on the PFD between Full Compass format and Arc format.− In Arc format the Weather Radar Display is also presented

whenever the Weather Radar is operating.

2 - GROUND SPEED AND TIME-TO-GO SELECTOR BUTTON− Pressing the GSPD/TTG Button alternates the respective

information on the PFD between ground speed and time-to-go.

3 - ELAPSED TIME SELECTOR BUTTON − The first actuation enters the Elapsed Time Mode on the PFD

respective field. The subsequent actuation provides thefollowing sequence of control: RESET - ELAPSED TIME -STOP - REPEAT.

4 - NAVIGATION SOURCES SELECTOR BUTTON − Provides the selection of the VHF NAV (VOR, ILS and MLS) as

navigation source for the EHSI. If the VHF NAV is alreadyselected, pressing the NAV Button selects the opposite VHFNAV as navigation source for the on-side EHSI. Pressing theNAV Button once again will restore the normal operation: VHFNAV 1 information presented on the PFD 1 and VHF NAV 2information presented on the PFD 2.

5 - FMS SOURCE SELECTOR BUTTON (OPTIONAL) − Provides the selection of the FMS as navigation source for the

EHSI.− On airplanes equipped with dual FMS, pressing the FMS Button

for the second time selects the opposite FMS as navigationsource for the on-side EHSI (and for the on-side MFD MAP).Pressing the FMS Button once again will restore the normaloperation: FMS 1 information presented on the PFD 1 (and MFD1) and FMS 2 information presented on the PFD 2 (and MFD 2).

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6 - BEARING SELECTOR KNOBOFF − The associated PFD bearing pointers are disabled.NAV 1 (2) − Selects the respective VHF NAV as source for the

associated bearing pointer.ADF − Selects the respective ADF as source for the

associated bearing pointer.FMS − Selects the FMS as source for the associated

bearing pointer.

7- DECISION HEIGHT SETTING AND IC-600 TEST KNOB − Provides the Radio Altimeter (RA) decision height setting.− When pressed on ground provides the IC-600 and RA test

activation. Refer to Section 2-4 – Crew Awareness for furtherinformation on test function and Section 2-17 – FlightInstruments for further information on decision height settingand RA test in flight.


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DISPLAYS CONTROL PANEL


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FMS SOURCE SELECTION ON THE MFD

As explained on the Display Control Panel (DCP) description, pressingthe FMS Button on that panel selects the FMS as navigation source forthe PFD and MFD.

On airplanes equipped with dual FMS, pressing the FMS Button (onthe Display Control Panel) for the second time selects the oppositeside FMS as navigation source for the on-side EHSI (and for the on-side MFD MAP). Pressing the FMS Button once again will restore thenormal operation: FMS 1 information presented on the PFD 1 (andMFD 1) and FMS 2 information presented on the PFD 2 (and MFD 2).

However, on airplanes equipped with dual FMS it is possible to selectthe opposite side FMS as MFD navigation source even if the FMS isnot selected as navigation source for the PFD.

In this case, pressing the MFD Bezel Button adjacent to the MFD SRClabel (presented on the MFD submenu), the on-side MFD will displaythe opposite side FMS data. This label is not presented if the FMS isalready selected as navigation source for the PFD.

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CROSS-SIDE FMS SOURCE SELECTION ON THE MFD


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ADF, VHF NAV AND DME INDICATIONS ON THE PFD

1 - VERTICAL DEVIATION DISPLAY− Color:

− Scale: white.− GS Pointer: - green.

- yellow if the same source is supplying bothsides.

− GS label: green.− For glide slope presentation the pointer will be parked up or

down of the deviation display when the deviation exceeds theexternal dots.

− Glide slope information will be displayed when SRN NAV isselected for display and tuned to LOC is active.

2 - MARKER BEACON DISPLAY− Color:

− O label: cyan.− M label: amber.− I label: white.− Box: white.

− An O, an M or an I flashing annunciation is displayed when theouter marker, the middle marker or the inner marker is detected,respectively.

− A beacon box surrounding the MB flashing annunciations will beshown when a SRN is displayed, tuned-to-localizer is active anda marker is also active.

3 - BEARING POINTER− Color:

− Cyan for Bearing 1− White for Bearing 2

− Circle coded for #1 source {VOR 1, ADF (for single installation)or ADF 1 (for dual installation)}.

− Diamond coded for #2 source {VOR 2, ADF (for singleinstallation) or ADF 2 (for dual installation)}.

− Pointer is removed if the selected source signal is invalid.

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4 - TO/FROM POINTER− Color: White.− Displayed towards the nose or the tail of the airplane to indicate,

respectively, "TO" or "FROM" the navigation aid.

5 - DME FIELD− Displays Ground Speed, Time-to-go, and Elapsed Time.− GROUND SPEED DISPLAY

− Color: Digits: green.GSPD label: white.

− Range: 0 to 550 KIAS.− Resolution: 1 KIAS.

− TIME TO GO DISPLAY− Color: Digits: the same of the NAV source color.

TTG label: white.− Range: 0 to 399 min.− Resolution: 1 minute.

− ELAPSED TIME− Color: Digits: green.

ET label: green.− Range: 00:00 to 09:59 h.− Resolution: Displayed in the format minutes: seconds (for

less than one hour), and hours (minutes for more than onehour).

6 - COURSE DEVIATION SCALE− Color: White.

7 - COURSE DEVIATION BAR− Color:

− Green: when the source is the on-side VOR.− Yellow: when the source is the cross-side VOR.

− Indicates against the course deviation scale, the differencebetween the selected course and the VOR bearing.


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8 - BEARING SOURCE ANNUNCIATIONS− Label: VOR1, VOR2, ADF1 or ADF2.− Color:

− Cyan for Bearing 1− White for Bearing 2

− Circle coded for #1 source {VOR 1, ADF (for single installation)or ADF 1 (for dual installation)}.

− Diamond coded for #2 source {VOR 2, ADF (for singleinstallation) or ADF 2 (for dual installation)}.

− Indicates the current source of input to the bearing pointers.− Source annunciation will be retained on the PFD, even in case

of an invalid bearing signal.

9 - DME HOLDING AND DISTANCE ANNUNCIATION− Color:

− Digits: green.− NM label: white.− H label: amber.

− Range:− Short Range NAV: 0 to 300 NM.

− Resolution: 0.1 NM.− When the DME hold is active an H label is displayed on the RH

of the DME distance digital readout. In this condition the H labelreplaces the distance NM label.

10 - COURSE DEVIATION NAV SOURCE ANNUNCIATION− Label: VOR1, VOR2, ILS1, ILS2 or FMS (optional)– Color:

– Yellow: when the same source is selected for both sides oris supplying cross-side.

– Green: when both sides present on-side sources, even ifthey are different.

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ADF, VHF NAV AND DME INDICATIONS ON THE PFD(EHSI IN FULL COMPASS FORMAT)


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ADF, VHF NAV AND DME INDICATIONS ON THE PFD(EHSI IN ARC FORMAT)


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FMS INDICATION ON THE PFD

1 - VERTICAL ALERT ANNUNCIATION− Label: VTA− Color: Amber− The VTA is displayed when the vertical alert bit is received from

the FMS.

2 - VERTICAL DEVIATION DISPLAY− When the FMS VNAV is selected the Vertical Deviation is

activated.− The Vertical Deviation Display indicates the vertical deviation

between the airplane and the selected vertical path.− Label: FMS− Color: Amber

− The FMS label and the scale are white.− If the FMS is the navigation source for only one side, the

pointer will be magenta, otherwise it will be amber.

3 - MESSAGE ANNUNCIATION− Label: MSG− Color: Amber− The MSG is displayed when a message is available on the FMS

Panel.

4 - GROUND SPEED/TIME TO GO DATA− Label: GSPD for Ground Speed. TTG for Time To Go.− Color: Labels and units are white.

− For single configuration, if the FMS is the navigation sourcefor only one side, the GSPD and TTG readouts will bemagenta, otherwise, they will be amber.

− For dual configuration, if each FMS is the navigation sourceof the respective side, the GSPD and TTG readouts will bemagenta. Otherwise, they will be amber.

− The Ground Speed unit is knots (KTS) and the Time To Go unitis minutes (MIN).

− The resolution of the digital values is 1 unit.− For invalid values, the digits will be replaced with three amber

dashes.

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5 - DRIFT ANGLE BUG− Color: Magenta.− The Drift Angle Bug rotates around the compass card, providing

the reading of the airplane tracking.

6 - COURSE DEVIATION BAR− Color: If the FMS is the navigation source for only one side, the

Course Deviation Bar will be magenta, otherwise, it will beamber.

7 - TO/FROM POINTER− Color: White.

8 - BEARING POINTER− Color: Cyan for Bearing 1 (circle shaped).

White for Bearing 2 (diamond shaped).

9 - BEARING SOURCE ANNUNCIATIONS− Color: Cyan for Bearing 1 (circle shaped).

White for Bearing 2 (diamond shaped) in single FMSconfiguration.In dual configuration there will be an indication if FMS 1or 2 is being used.

10 - WIND VECTOR DISPLAY− Color: Magenta.− A single vector shows the direction of the wind relative to the

airplane symbol. The digits indicate the wind intensity in knots.

11 - DEGRADE MODE/DEAD RECKONING MODE/WAYPOINTANNUNCIATIONS− Label: DGRAD for Degrade Mode (single FMS configuration only)

DR for Dead Reckoning mode.WPT for waypoint.

− Color: Amber− WPT is lit when the airplane is approaching the next waypoint.

12 - DISTANCE DISPLAY− Color:

− In single configuration, if the FMS is the navigation sourcefor only one side, the distance readout will be magenta.Otherwise, it will be amber.

− In dual configuration, if each FMS is the navigation sourceof the respective side, the distance readout will bemagenta, otherwise it will be amber.

− The unit is white.− The distance unit is nautical miles (NM).


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13 - TO WAYPOINT SYMBOL− Label: Waypoint identifier name (Ex: KDVT).− Color: Magenta. For dual configuration, when using cross-side

information, the color is amber.− In the sequence established, the TO waypoint is the next one

from the current airplane position.

14 - APPROACH/TERMINAL AREA ANNUNCIATIONS− Label: APP for Approach.

TERM for Terminal Area.− Color: Cyan.− When APP is displayed it indicates that the FMS is in the flight

approach phase and also can indicate that the lateral deviationscaling has been set to approach scale factor.

− In the APP mode the deviation indicator sensitivity and FMStracking gains are increased.

− The TERM annunciator is displayed when the airplane entersin the terminal area or when the lateral deviation scaling hasbeen set to the enroute scale factor.

− Priority is given to the APP message.

15 - FMS SOURCE ANNUNCIATION− Label: FMS.− Color:

− For single configuration, if the FMS is the navigation sourcefor only one side, the FMS label will be magenta.Otherwise, it will be amber.

− For dual configuration, if each FMS is the navigation sourcefor the respective side, the FMS label will be magenta,otherwise it will be amber.

− FMS is displayed only when a single source is installed.

16 - HEADING ANNUNCIATION− Label: HDG SEL (For dual FMS configuration).− Color: White. For dual configuration, if each FMS is the

navigation source for the respective side the color will be white, otherwise it will be amber.

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17 - SELECTED COURSE/DESIRED TRACK ANNUNCIATIONS ANDREADOUTS− Label: DTK for Desired Track.

CRS for Selected Course.− Color:

− For single configuration, if the FMS is the navigation sourcefor only one side, the CRS label will be green and DTK willbe magenta. Otherwise, both labels will be amber.

− For dual configuration, if each FMS is the navigation sourcefor the respective side, the CRS and DTK labels will bemagenta. Otherwise they will be amber.

− The readouts will have the same color as the CRS and DTKannunciations.

− DTK is displayed when the FMS is the selected navigationsource.

18 - CROSSTRACK ANNUNCIATION− Label: SXTK− Color:

− For single configuration, if the FMS is the navigation sourcefor only one side the label will be magenta, otherwise it willbe amber.

− For dual configuration: The color will be ever amber.

− SXTK is displayed to indicate that the airplane is off track.

19 - CAPTURED LATERAL MODE− Refer to Section 2-19 - Autopilot.


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FMS INDICATION ON THE PFD


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FMS INDICATION ON THE MFD

1 - FMS SOURCE ANNUNCIATION− Label: FMS for single configuration.

FMS1 or FMS2 for dual configuration.− Color:

− Magenta: when the source is the on-side FMS.− Yellow: when the source is the cross-side FMS.

2 - DRIFT ANGLE BUG− Color:


− The Drift Angle Bug rotates around the compass card, providingthe reading of the airplane tracking.

3 - WAYPOINT SYMBOL− Label: Waypoint identifier name (Ex: KDVT).− Color: All Waypoints are white except the TO waypoint.− Waypoint is displayed as a four pointed star at the geographical

locations, referenced to the current present position, where theselected transitions of the flight plan occur.

− A maximum of 10 Waypoints can be displayed, including theFROM waypoint.

− A navigation aid or airport can also be located on the flight planat a transition point and is accounted in the maximum allowablenumber of Waypoints.

4 - AIRPORT ANNUNCIATION− Label: APT.− Color: Cyan.− Appears when an airport symbol is shown along the route.

5 - NAVAID ANNUNCIATION− Label: NAV.− Color: Cyan for single or green for dual configuration.− Appears when a navaid symbol is shown along the route.

6 - DESIGNATOR RANGE AND BEARING READOUT− Color: Cyan.− The range readout indicates the distance between the airplane

and the Designator Symbol.− The bearing readout bearing location of the Designator Symbol

related to the airplane position.

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7 - TO WAYPOINT SYMBOL− Color:


− In the sequence established, the TO Waypoint is the next onefrom the current airplane position.

8 - LATERAL DEVIATION DISPLAY− Color: White.− Right after the values there is a letter which may be L or R

standing for Left and Right respectively.

9 - WIND VECTOR DISPLAY− Color:


− A single vector shows the direction of the wind relative to theairplane symbol. The digits indicate the wind intensity in knots.

10 - DESIGNATOR SYMBOL− Color:

− Same color of the Waypoint: If the Designator is co-locatedwith a connected Waypoint.

− Cyan: If it is not connected.− The Designator symbol is displayed as an unfilled rectangle

applied in two distinct methods: co-located with a Waypoint orpositioned with the joystick.

− Designator will not be displayed if it represents the currentposition.

11 - TO WAYPOINT DATA ANNUNCIATIONS− It is composed of the annunciators and presented as follows:

− Identification.− Distance in nautical miles (NM).− Time to the TO Waypoint in minutes (MIN).

− Color:− For single FMS configuration the identification is magenta.

The distance and the time are white.− For dual FMS configuration the identification, distance and

time are magenta, when the source is the on-side FMS, oryellow, when the source is the cross-side FMS.


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FMS INDICATION ON THE MFD


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WEATHER RADAR SYSTEMThe airplane can be equipped with P-660 or P-880 weather radarmodels and 12 inch antenna. For additional information on functionsand operations, refer to the manufacturer’s manual.

The weather radar system is designed for detection and analysis ofprecipitation in storms along the flight path of the airplane. The systemprovides the flight crew with visual indications regarding rainfallintensity and turbulence content.

Precipitation intensity level is displayed in four bright colors (magenta,red, yellow and green) contrasted against a deep black background onthe PFDs’ and MFDs’ radar mode field. Magenta represents theheaviest rainfall intensity while green indicates the lightest.

The radar may also be used for ground mapping. When operating inground mapping mode, prominent landmarks are displayed, which allowsidentification of coastlines, mountainous regions, cities or even largestructures.


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GENERAL

The weather radar system consists of an integratedReceiver/Transmitter/Antenna unit (RTA) and a dedicated controlpanel. The RTA transmits and receives on the X-band radio frequency.The RTA processes radar echoes received by the antenna. The scan-converted data are displayed on PFDs’ and MFDs’ radar mode field.

The weather radar system run on 28 V DC powered by one of theAvionics Switched DC Buses. Should a power supply failure occur, theweather radar system will become inoperative, as there is no backuppower source for this system.

The weather radar interfaces with other airplane systems andequipment as presented in the schematic diagram below:

WEATHER RADAR SCHEMATIC


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WEATHER RADAR NORMAL OPERATION

The weather radar is controlled through the weather radar controlpanel and via the MFD Bezel Buttons. The weather radar control panelprovides control functions and operating modes management forproper weather radar operation. The airplane is equipped with twoweather radar control panels located on the glareshield panel.

INTERPRETING WEATHER RADAR IMAGES

The weather radar is a water detector. It is calibrated to best see waterin its liquid form and with an ideal raindrop diameter. The weatherradar can see rain, wet snow, wet hail and dry hail (depending on itsdiameter). The radar can not see water vapor, ice crystals and smalldry hail.

At higher altitudes, there is less humidity in the air and consequentlythere is less water condensation. It means that heavy precipitation anddense cells are less likely to occur. As a result, flight level 200 (FL200)is defined as "FREEZING LEVEL", i.e., presence of water in its liquidform is not forecast above this level. However, CBs and otherphenomena may push humidity and water, sometimes supercooledwater, to higher altitudes due to convective activity.

WARNING: DRY HAIL CAN BE PREVALENT AT HIGHERALTITUDES. SINCE ITS RADAR REFLECTIVERETURN IS POOR, IT MAY NOT BE DETECTED.

Use increased gain when flying near storm tops in order to display thenormally weaker returns that could be associated with hail.

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RADAR WARM UP PERIOD

When power is first applied to the radar, a period of 40 to 100 secondsis required to allow its magnetron to warm up. The radar displays theWAIT message on the PFDs’ and MFDs’ radar mode field and doesnot transmit or perform an antenna scan. After the completion ofwarm-up period, the radar automatically become operational in theselected mode or goes to forced standby (FSBY) if the airplane is onthe ground.

GROUND OPERATION PRECAUTIONS

If the radar system is to be operated in any mode other than standby orforced standby while the airplane is on the ground, the followingprecautions should be taken:

- Direct nose of airplane so that antenna scan sector is free of largemetallic objects such as hangars or other airplanes for a distance of30 meters (100 ft).The antenna must be tilted fully upwards.

- Avoid using the weather radar during airplane refueling or within 30meters (100 ft) of any other airplane undergoing refueling operations.

- Avoid using the weather radar if personnel are standing too close tothe 270° forward sector of airplane.

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WEATHER RADAR OPERATING MODES ANDFUNCTIONS

TEST MODE (TST)

After the radar warm-up period is over, the TEST mode may beselected. A special test pattern made up of color bands is displayed. Aseries of green/yellow/red/magenta/white bands indicate that the signalto color conversion circuits are operating normally. A 100 mile range isautomatically selected. A green TEST label will be displayed on thePFDs’ and MFDs’ radar mode field.

When the airplane is on the ground and the TEST mode is entered, thefirst page always includes RADAR OK or RADAR FAIL to indicate thecurrent state of the radar, as follows:RADAR OK: indicates that no faults were found and the radar is readyfor service. It is combined with the END OF LIST page.RADAR FAIL: indicates a radar fault.

During the weather radar test, several fault messages may bepresented to the crew. The POC (Power On Counter), aside recordingan existing fault, also stores fault information from previous power-oncycles. However, if the first page announces "RADAR OK", the radar isready for service.

STANDBY MODE (SBY)

The standby mode should be selected any time it is desired to keepthe system powered without transmitting. When SBY mode is selectedthe WX radar remains in a ready state, with the antenna scanmotionless and stowed in a tilt-up position. In addition, the transmitteris inhibited and the display memory is erased.

Placing only one controller in SBY does not shut the transmitter OFF.Instead, the no-SBY controller governs radar operation. If bothcontrollers are placed in SBY, the transmitter is shut OFF.

In standby mode a STBY label is displayed on the PFDs’ and MFDs’radar mode field.


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FORCED STANDBY MODE (FSBY)

The FSBY is an automatic, non-selectable radar mode, that forces theradar into standby when the airplane is on the ground (weight-on-wheels logic) regardless of the selected active radar mode. This is asafety feature that inhibits the transmitter on the ground to eliminateX-band microwave radiation hazards. In FSBY mode, the transmitterand the antenna scan are both inhibited, memory is erased and aFSBY label is displayed on the PFDs’ and MFDs’ radar mode field.

The forced standby mode may be overridden on the ground by pushingthe STAB button 4 times in 3 seconds.

CAUTION: IF FSBY MODE IS OVERRIDEN ON THE GROUND ANDANY RADAR ACTIVE MODE IS SELECTED, THETRANSMITTER IS TURNED ON. THE RADAR MUSTNOT BE OPERATED UNDER THIS CONDITION WHILEREFUELING, NEAR FUEL SPILLS OR PEOPLE.

WEATHER DETECTION MODE (WX)

The WX mode is used to detect areas of severe weather. This willallow the pilots to avoid dangerous weather conditions and possibleturbulence areas. WX may be used on the ground, often prior totakeoff, in order to monitor the weather in the immediate vicinity. In thiscase, the forced standby mode may be overridden.

In WX Mode, the weather radar system is fully operational and allinternal parameters are set for enroute weather detection. A WX labelis displayed on the PFDs’ and MFDs’ radar mode field.

The levels and colors associated with the storm category are asfollows:

LEVEL COLOR RAINFALL CATEGORY4 Magenta Extreme/Intense3 Red Very Strong/Strong2 Amber Moderate1 Green Moderate/Weak0 Black Weak

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RAIN ECHO ATTENUATION COMPENSATION TECHNIQUEFUNCTION (REACT or RCT)

The REACT is a sub mode of the weather detection mode and whenselected activates three separate but related functions:

− Attenuation Compensation - Storms with high rainfall rates canattenuate the radar energy making it impossible to see a second cellhidden behind the first cell. In the REACT mode, the radar incorporates a function thatautomatically adjusts receiver gain by an amount equal to the amountof attenuation, i.e., the greater the amount of attenuation, the higherthe receiver gain and thus, the more sensitive the receiver.

− Cyan REACT Field - Since there is a maximum limit to receiver gain,

strong targets (high attenuation levels) cause the receiver to reach itsmaximum gain value and weather targets can no longer be calibrated.The point where red level weather target calibration is no longerpossible is highlighted by changing the background field from black tocyan. Cyan areas should be avoided. Any target detected inside a cyanarea should be considered very dangerous. All targets in the cyanfield are displayed as a magenta-colored 4th level precipitation.

− Shadowing - This is an operating technique similar to the CyanREACT Field. To use the shadowing technique, tilt the antenna downuntil the ground is being painted just in front of the storm cell(s). Anarea characterized by no ground returns behind the storm cell has theappearance of a shadow. The cell that produces radar shadowing is avery strong and dangerous cell and should be avoided by 20 NM.

FLIGHT PLAN MODE (FP)

When the Flight Plan Mode is selected a singular display of navigationdata and a FLTPLAN label are presented on the PFDs’ and MFDs’radar mode field. The radar is put in standby and there is no radar datadisplayed in this mode.


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GROUND MAPPING MODE (GMAP)

This mode is used to alert the flight crew regarding hazards caused byground targets. This is especially useful in areas of rapidly changingterrain, such as mountainous regions. In this mode the system is fullyoperational and all internal parameters are set to enhance returns fromground targets.

The TILT control should be turned down until the desired amount ofterrain is displayed. The degree of down-tilt depends upon airplanealtitude and the selected range. Receiver characteristics are altered toprovide equalization of ground-target reflection versus range. Theselection of calibrated GAIN will generally provide the desired mappingdisplay. If required, variable gain may be used to reduce the level ofstrong returns.

In the ground mapping mode a GMAP label is displayed on the PFDs’and MFDs’ radar mode field, and the color scheme is changed to cyan,yellow and magenta. Cyan represents the least reflective return, yellowis a moderate return and magenta represents the most highly reflectivetarget return.

It is possible to have one pilot working the GMAP while the other one isusing the regular WX mode.

CAUTION: WEATHER TYPE TARGETS ARE NOT CALIBRATEDWHEN THE RADAR IS IN THE GMAP MODE.THEREFORE, THE PILOT SHOULD NOT USE THEGMAP MODE FOR WEATHER DETECTION.

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TURBULENCE DETECTION FUNCTION (TRB) (P-880 MODELONLY)

When this mode is selected, the radar processes return signals inorder to determine if a turbulence condition is present. Areas ofpotentially hazardous turbulence are displayed as white. Any areasshown as turbulence should be avoided.

Turbulence detection function may only be engaged in the WX modeand at selected ranges of 50 NM or less. When the TRB function isactive, a T letter will be displayed on the PFDs’ and MFDs’ radar modefield.

CAUTION: ALTHOUGH TURBULENCE MAY EXIST WITHIN ANYSTORM CELL, WEATHER RADAR CAN ONLY DETECTTURBULENCE IN AREAS OF RAINFALL.

TARGET ALERT (TGT)

Target alert is selectable in all but the 300 mile range. When selected,target alert monitors for red or magenta weather beyond the selectedrange and 7.5° on either side of the airplane’s heading. If such weatheris detected within the monitored area and outside the selected range,the target alert annunciation TGT label changes from a green armedcondition to an yellow TGT alert condition on the PFDs’ and MFDs’radar mode field. This annunciation advises the flight crew thatpotentially hazardous targets lie directly in front and outside of theselected range. When this warning is received, the flight crew shouldselect longer ranges to view the questionable target.

The target alert is inactive within the selected range. Selecting targetalert forces the system to calibrate gain, and turns off the variable gainmode. Target alert can only be selected in WX and FP modes.

NOTE: Keep TGT alert enabled when using short ranges. This allowsthe issuing of an alert if a new storm cell develops ahead of theairplane’s flightpath.


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ANTENNA STABILIZATION (STAB or STB)

The antenna is normally pitch and roll-stabilized by using attitudeinformation from the IRS. Momentarily pushing the STAB (or STB)button disables antenna stabilization and an amber “STAB”annunciation label is presented on the PFDs’ and MFDs’ radar modefield.

RECEIVER GAIN (GAIN)

The GAIN knob is a rotary control and push/pull switch that controlsradar receiver gain. Two gain modes are available: calibrated orvariable.

Calibrated: When the GAIN knob is pushed in, receiver gain is presetand calibrated, which is the normal mode of operation. In calibratedgain, the rotary function of the GAIN knob is disabled.

Variable (VAR): When the GAIN knob is pulled out, the system entersthe variable gain mode. Variable gain is used for additional weatheranalysis and for ground mapping. In the WX mode, variable gain canincrease receiver sensitivity over the calibrated level to show very weaktargets or can be reduced below the calibrated level to eliminate weakreturns. In the GMAP mode, variable gain is used to reduce the level ofstrong returns from ground targets.

Rotation of the knob counter-clockwise reduces receiver sensitivity.Rotating clockwise increases receiver sensitivity until its maximum. Adigital readout and gain setting label are displayed on the PFDs’ andMFDs’ radar mode field.

NOTE: When REACT or TGT modes are selected, the system will beforced into calibrated gain.

CAUTION: VARIABLE GAIN MAY BE USED ONLY FOR SHORTPERIODS OF TIME. DO NOT LEAVE THE RADAR INVARIABLE GAIN SINCE SIGNIFICANT WEATHERTARGETS MAY NOT BE DISPLAYED.

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TILT

Tilt management is crucial to the safe operation of weather radar. Ifimproperly managed, weather targets can be missed orunderestimated. Proper tilt management demands that tilt be changedcontinuously.

To find the best tilt angle after the airplane is airborne, adjust the TILTantenna downward until a few ground targets are visible at the edge ofthe display. The table below gives the approximate tilt settings forminimal ground target display for different altitudes and ranges. If thealtitude changes or a different range is selected, adjust the tilt controlas required to minimize ground returns.

When flying at high altitudes, tilt downward frequently to avoid flyingabove storm tops. When in low altitude or approaching for landing, tiltmanagement must be performed manually, with the radar beamvertically sweeping from up to down to avoid flying above or below astorm line.

During takeoff, the radar must be adjusted to a minimum range scale,with a horizontal RH and LH scan and with the antenna positionedupwards (climbing angle).

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TILT SETTINGS FOR MINIMAL GROUND TARGET DISPLAY(12 inch antenna)


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The figure below helps to visualize the relationship between tilt angle,flight altitude and selected range. It shows the distance above andbelow airplane altitude that is illuminated by the radar during level flightwith 0° tilt (high altitude) and a low altitude situation, with antennaadjusted for 2.8° up-tilt.

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ALTITUDE COMPENSATED TILT (ACT) (P-880 MODEL ONLY)

In ACT, the antenna tilt is automatically adjusted with regard to theselected range and airplane altitude. ACT adjusts the tilt to show a fewground targets at the edge of the display. The TILT knob can be usedfor fixed offset corrections of up to 2°.

NOTE: Proper tilt management demands that tilt be changedcontinuously, even in airplanes equipped with ACT.

SLAVE (SLV)

One controller can be slaved to the other by selecting OFF on thatcontroller only. This condition is annunciated by the illumination of SLVon the control panel. The slave mode allows one controller to set theradar modes for both sides. In the slave mode, the PFDs and MFDsradar information are identical and simultaneously updated.

NOTE: In the slaved condition, both control panels must be set to offbefore the radar system turns off.


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RADOME

The radome is the primary factor behind degraded weather radarperformance. The problems affecting the radome are as follows:- A water film over the radome’s surface when flying in rain.

- Greased radome.

- Cracked radome.

- Holes caused by lightning strike/electrostatic discharges.

- Excessive application of antistatic paint.

Water Film Over The Radome’s Surface: When flying in rain, thereis indication that at some specific altitudes and speeds a water film isformed on the radome, altering the weather radar indications. Theradar display may disappear or turn red. To avoid this problem, there isa hydrophobic coating product named Cytonix that can be applied tothe radome surface.

Greased Radome: The presence of grease or dirt over the radome’ssurface also impairs radar transmission. These should be reportedimmediately to maintenance personnel for cleaning or correctiveaction.

Electrostatic Discharges: Static electricity influences radarperformance. The right bonding is necessary. Bonding is accomplishedthrough two metallic meshes that link the radome’s metallic bulkhead(diverters) to the airplane’s airframe. It is important to make sure thatthey are in good condition and not painted. If both the metallic meshesand screws are painted, this will isolate the static power generated inthe radome, resulting in electrical discharges that will follow towardsthe radar antenna and/or generate noise in the audio system.

Cracked Radome: Small holes caused by electrostatic discharges,minor damage to structure or paint can cause water infiltration in theradome’s honeycomb composite structure. It can result in significantradar signal attenuation, distortion and in some cases, can cause darkspots on the radar screen.

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WEATHER RADAR CONTROLS AND INDICATORS

WEATHER RADAR CONTROL PANEL

1 - RANGE SELECT BUTTONS − Allow selection of the radar’s operating range, from 5 to 300 NM

full scale in WX, REACT, or GMAP mode. In FP mode,additional ranges of 500 and 1000 NM are available. In testmode the range is automatically set to 100 NM.

− The up-arrow button selects increasing ranges, while thedown-arrow button selects decreasing ranges. Upon reachingmaximum or minimum range, further pushing of the buttoncauses the range to rollover to minimum or maximum range,respectively.

2 - TURBULENCE DETECTION FUNCTION BUTTON (P-880 ModelOnly)− Alternate pressings turns on or off the radar’s turbulence

detection function.− Function can be used only in WX or RCT mode, with selected

range of 50 NM or less.

3 - STABILIZATION FUNCTION BUTTON− When momentarily pressed, disables antenna stabilization

function. The STAB OFF annunciator will illuminate on thecontrol panel.

− On the ground, after warm-up period, pressing the STB buttonfour times within 3 seconds will inhibit the forced standby(FSBY) function.

4 - SLAVE ANNUNCIATOR− Illuminates to indicate that one controller is slaved to the other.

5 - TARGET ALERT CONTROL BUTTON− Alternate pressing selects or cancels the target alert feature.− Selectable only in the WX and FP Modes.

6 - SECTOR SCAN BUTTON (SECT)− When momentarily pressed, selects either the radar’s normal 12

sweeps per minute for a 120° full scan or the faster update 24sweeps per minute for a 60° sector scan.


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7 - ANTENNA TILT CONTROL KNOB− The TILT knob is a rotary control that allows manual control of

the antenna’s tilt angle. Clockwise rotation tilts the beam upward0° to +15°. Counter-clockwise rotation tilts beam downward 0° to–15°. A digital readout of the antenna tilt angle is displayed onthe MFD.

− The range between +5° and -5° is expanded for setting ease.

ALTITUDE COMPENSATED TILT (PULL ACT) (P-880 Model Only)− Pulling out the TILT knob activates the auto tilt control, which

automatically readjusts tilt between ± 2° based on changes inbarometric altitude and/or selected range.

8 - RADAR MODES CONTROL KNOB OFF- Turns off the weather radar. SBY - Selects the weather radar standby operating mode. WX - Selects the weather radar detection operating mode. RCT- Selects the REACT function (P-880 Model only). GMAP - Selects the weather radar ground mapping operatingmode. FP - Selects the weather radar flight plan operating mode. TST - Selects the weather radar test mode.

9 - GAIN CONTROL KNOB− Allows receiver gain control.− When pushed in, receiver gain is preset and calibrated. Rotary

function of the GAIN knob is disabled.− When pulled out, sets receiver gain to variable (VAR) mode.

10 - RAIN ECHO ATTENUATION COMPENSATION TECHNIQUEFUNCTION BUTTON (P-660 Model Only)

− When pressed (momentarily), enables the REACT.− REACT is always selected in test mode.− REACT is available in all modes except MAP.

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WEATHER RADAR CONTROL PANEL


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MFD BEZEL PANEL

1 - WEATHER RADAR DISPLAY SELECTOR BUTTON− − Alternate pressing of the weather radar display selector button

allows the weather radar to be displayed or removed from theMFD. Control of all other weather radar functions isaccomplished by the radar control panel. When the weatherradar is selected, the WX label on the MFD menu, above thisbutton, will be highlighted by a white box.

− − The weather radar can only be selected for display in mapformat. If the weather radar is selected with plan format alreadyselected on the MFD, it will force the display to revert to mapformat.

2 - MAP/PLAN FORMATS CONTROL BUTTON− Alternate pressing of the map/plan formats control button will

cause the MFD to toggle between map and plan formats. Awhite box around will highlight the selected MFD format.

− If the weather radar is displayed on the MFD and the plan formatis selected, the weather radar will be removed from the display.However, if the MFD map format is selected again, the weatherradar display will be restored on the MFD.

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MFD BEZEL PANEL


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WEATHER RADAR DISPLAY ON THE PFD AND MFD

1 - ANTENNA POSITION INDICATOR (API)− Color: Amber.− The API is displayed as an arc at the current range outer limit.− Indicates the radar antenna alternate sweep position and

provides a picture bus activity indication.

2 - WEATHER RADAR PATCH− Indicates an area of radar reflection.− Color:

− Magenta: high intensity reflection.− Red: medium-high intensity reflection.− Yellow: medium intensity reflection.− Green: low intensity reflection.

3 - WEATHER RADAR TURBULENCE INDICATION− Indicates an area of detected turbulence.− Color: white.

4 - WEATHER RADAR REACT INDICATION− Indicates an area where radar receiver gain compensation has

reached its maximum value.− Color: cyan.

5 - WEATHER RADAR RANGE ARC VALUE− Color: white.− Indicates the radar range selected in the weather radar control

panel.

6 - WEATHER RADAR ANTENNA TILT ANGLE DISPLAY− Color: green.− Range: –15 to +15°.− Resolution: 1°.

7 - WEATHER RADAR TARGET MODE AND ALERT ANNUNCIATION− Color:

− TGT label: green or amber.− VAR label: amber.

− The VAR label will be displayed in the same field as that usedfor TGT annunciation to indicate a variable gain indication.Priority is given to TGT annunciation.

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8 - WEATHER RADAR MODES ANNUNCIATION DISPLAY− Indicates the selected mode in the weather radar control panel.

DISPLAY MODE DESCRIPTION

ANNUNCIATION COLOR

STAB AMBER Stabilization off.TGT GREEN Target alert enable.TGT AMBER Target alert enable and level 3

WX return detected in the forward15° of antenna scan.

VAR AMBER Variable gain.WX GREEN Normal WX ON and selected for

display.WX AMBER Invalid WX control bus.TX GREEN WX is transmitting but not

selected for display, or in STBYor FSTBY.

TX AMBER WX is transmitting and weight onwheels indicates on ground, butnot selected for display, or inSTBY and FSTBY.

WAIT GREEN Warm up period of approximately40 to 100 seconds.

STBY GREEN Normal standby.FSBY GREEN Forced standby.TEST GREEN Test mode and no faults.FAIL AMBER Test mode and faults.RCT GREEN Normal WX with REACT.FPLN GREEN Flight plan mode.GMAP GREEN Ground map mode.GCR AMBER Normal WX with ground clutter

reduction.R/T GREEN WX with REACT and turbulence.WX/T GREEN Normal WX with turbulence.


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WEATHER RADAR DISPLAY ON PFD

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WEATHER RADAR DISPLAY ON MFD


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LIGHTNING SENSOR SYSTEM (LSS)The P-880 Weather Radar Model may be optionally equipped with aLightning Sensor System (LSS). For additional information on functionsand operations, refer to the LSZ-850 Lightning Sensor System Pilot’sHandbook.

The Lightning Sensor System is designed to detect and locate areas oflightning activity, in a 100-nautical-mile radius around the aircraft, andto give the operator a visual display of its position and rate-of-occurrence on the MFDs.

The Lightning Sensor System is inhibited on the ground and during HFtransmission (HF PTT), being powered by one of the 28 V DC AvionicsSwitched DC Buses. Should a power supply failure occur, the LightningSensor System will be inoperative, as there is no backup power to it.

The Lightning Sensor System interfaces with other airplane systemsand equipment as follows:

− MFDs - The LSS provides lightning activity data to the MFDs.− IC-600s - Signals between the LSS and the IC-600s are transmitted

through a Serial Data Bus.

LIGHTNING SENSOR SYSTEM SCHEMATIC


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LSS OPERATION

The Lightning Sensor System is controlled through a dedicated controlknob, located on the Weather Radar Control Panel. The LightningSensor System Control knob provides control functions and operatingmodes management for proper Lightning Sensor System operation.

The Lightning Sensor System detects both visible and invisible highenergy electromagnetic discharges (lightning), indicating areas ofturbulent activity and displaying such information on the MFDs.

The Lightning Sensor System may be operated with the WeatherRadar System turned on or off. If the Weather Radar is turned off, theLSS Range Selection is done through the MFDs knobs.

When power is first applied to the system, a self-test is automaticallyperformed and, in the event of failure, an amber LX/F will be displayedon the MFDs. The LSS also performs self-calibration each time thesystem is turned on. The green LX/C annunciation is removed aftercalibration is completed. Strong interfering signals outside the aircraft,or equipment malfunction may prevent the calibration of the system, inwhich case the ¨C¨ will not be removed from the display. If the ¨C¨ isremoved after takeoff, this means that only outside interference waspreventing calibration. In this case the LSS should be switched off andset back to LX mode to force recalibration for greater accuracy. If the¨C¨ persists after takeoff, a test should be performed.

LSS OPERATING MODES AND FUNCTIONS

STANDBY MODE (STBY)

When the Standby Mode is selected, no lightning data is shown on theMFDs. However, the receiving and processing equipment is active andlightning strikes are being counted and accumulated into areas.

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LIGHTNING DETECTION MODE (LX)

When operating in Lightning Detection Mode, the Lightning SensorSystem is fully operational and lightning strikes are collected,processed, and displayed.

The LSS shows areas of lightning activity with white lightning symbols,which represents the center of a circular area with a radius of ninenautical miles. Three different lightning symbols are used to representthree different lightning rates-of-occurrence within each 18-milediameter cycle.

Because of the large variation in lightning electromagnetic discharge,sometimes the system can not determine, within its occurring criteria,lightning bearing and distance. Rather, only the bearing is measured.Activity that only provides bearing information is presented in magentalightning symbols near the outer range marks.

When operating the LSS with the Radar in STBY Mode, 360º of data isdisplayed.

A green LX will be displayed on normal operation.

Lightning Sensor Display symbology is as follows:

LEVEL LIGHTNING CATEGORY1 white headless lightning symbol2 white single-headed lightning symbol3 white double-headed lightning symbol

ALERT magenta headless lightning symbol

CLEAR/TEST MODE (CLR/TST)

When the Clear/Test mode is selected, all lightning rate symbols areerased from the display and a special pattern is displayed to allowverification of the Lightning Sensor System operation.The Test must be accomplished by selecting 50 NM or a greaterdisplay range and the CLR/TST Mode. The LSS CLR/TST Mode canbe selected in any radar mode.In the event of a failure, an amber LXmn will be displayed, where mn isa failure code, which will help the technician in troubleshooting thesystem.When CLR/TST mode is selected, a green LX/CL is displayed on theMFDs and, after three seconds, LX/CL is replaced by LX/T.


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LSS CONTROLS AND INDICATORS

1- LIGHTNING SENSOR SYSTEM CONTROL KNOBOFF - Turns off the Lightning Sensor System.

SBY - Selects the Lightning Sensor System Standby Mode.

LX - Selects the Lightning Sensor System Detection OperatingMode.

CLR/TST- Selects the Lightning Sensor System Test Mode.

WEATHER RADAR CONTROL PANEL WITH LSS


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LIGHTNING SENSOR SYSTEM DISPLAY ON THE MFD

1 - LEVEL 1 LIGHTNING ACTIVITY INDICATION− Color: White− Indicates the center of an 18-mile diameter area of lightning

activity with 1 strike every 2 minutes.


activity with 2 strikes every 2 minutes.


activity with 3 strikes every 2 minutes.

4 - LEVEL 4 (ALERT) LIGHTNING ACTIVITY INDICATION− Color: Magenta− Indicates only the bearing of lightning activity, without

determining the distance.

5 - LIGHTNING SENSOR SYSTEM MODES AND FAULTSANNUNCIATION DISPLAY− Indicates, above the Weather Radar annunciation display, the

selected mode in the Lightning Sensor System Control Knob .

DISPLAY MODE DESCRIPTION

ANNUNCIATION COLOR

LX/OFF GREEN LX Power is off.LX AMBER LX Power is on and LX

Communication Bus is Inactive.LXmn AMBER LX Fault Code Enabled.LX/F AMBER LX FaultSTBY GREEN LX StandbyLX/CL GREEN LX ClearLX/T GREEN LX Test ModeLX/I GREEN LX Antenna InhibitLX/C GREEN LX Auto CalibrateLX GREEN LX Normal

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LIGHTNING SENSOR SYSTEM DISPLAY ON MFD


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IDENTIFICATION FRIEND OR FOE SYSTEM (IFF)The Mode-S/IFF Transponder System integrates military IFF functionsand civilian Mode-S/ATCRBS surveillance functions into a singlesystem.

The IFF system is a cooperative surveillance system designed toclassify airborne targets as friendly or hostile. It employs ground-basedand airborne interrogations and transponders.

A key feature of the Mode-S system that distinguishes it from ATCRBSis that each airplane is assigned a unique address code. Using thisunique code, interrogations can be directed to a particular airplane andreplies can be positively identified. Channel interference is minimizedbecause a sensor can limit its interrogations to targets of interests. Byproper timing of interrogations, replies from closely spaced airplanecan be received without mutual interference.

The Mode-S/IFF Transponder System consists of the following majorcomponents: one Mode-S transponder, one Mode-S/IFF transponder,Mode-S/IFF control panel, one selector panel and top and bottomantennas.

The transponder is capable of receiving/transmitting signals throughany one antenna, thus inhibiting the shadow effect caused by theairplane structure while maneuvering. The reply signal provides anidentity, an altitude, an identification position, or an emergencymessage. The Mode-S/IFF transponder operates in Mode 1, 2, 3/A, Cand S, the system receiving interrogations on 1030 MHz andtransmitting on 1090 MHz.

Avionics Master DC Bus 2A powers the Mode S/IFF transponder andthere is a 7.5 A circuit breaker for power wiring protection. The Mode-S/IFF Transponder System will not operate in the event of either afailure in the control panel or in the event of an electrical/generatorfailure.

The transponder reply-transmit capability is disabled when the systemis in STANDBY mode. The Weight-On-Wheels (WOW) circuitrydisables the ATCRBS’ transponder reply capability while the airplane ison the ground. However, the air/ground switch does not disable Mode-S transponder replies and airplane status is included in the Mode-S/IFFtransponder reply data.


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SELECTOR PANEL

The selector panel, on the main control panel, provides the means forthe flight crew to select either the Mode-S on transponder 1 or theMode-S/IFF transponder, according to mission requirements at thatparticular moment.

The Mode-S/IFF transponder will be used normally in the mission area.

1 - IFF POSITION

Selects the Mode-S/IFF transponder. All indications on the RMU pagesare related to the transponder 1 system and TCAS functions will bedashed. In this case all TCAS II operations is made through the Mode-S/IFF control panel.

2 - XPDR POSITION

Selects the Mode S transponder. Both Mode-S/IFF transponder unitand Mode-S/IFF control panel are turned off.

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IFF TRANSPONDER CONTROLS AND INDICATORS

The Mode-S/IFF control panel is located on the control pedestal andallows the selection of the Mode-S/IFF transponder operation modes(Off, Standby, Normal and Emergency), selection of interrogation andreply modes (1, 2, 3/A, C and S), and selection of codes.

1 - DISPLAY SELECT ROTARY KNOB

The Mode-S/IFF control panel has a six position rotary knob to selectthe different display modes.

MODE 1: Selecting “1” with the DPL SEL knob enables Mode 1display and keypad entry mode.

MODE 2: Selecting “2” with the DPL SEL knob enables Mode 2display and keypad entry mode.

MODE 3/A: Selecting “3A” with the DPL SEL knob enables theMode 3/A display and keypad entry mode.

MODE S: Selecting “S” with the DPL SEL knob enables theMode S display and keypad entry mode.

FLIGHT ID: Selecting “FLT ID” with the DPL SEL knob enablesFLIGHT ID display and keypad entry mode.

The Mode-S Aircraft Identification Subfield (AIS) allowsthe operator to select the airplane Flight ID to bedownlinked as part of Downlink Format (DF) message20 and 21 from the control panel. The Mode S FlightID is typically the airplane flight number or airplaneregistration number.

STAT: Messages from the transponder are displayed on theLCD only if the STAT position in the DPL SEL knob isselected. Transponder messages have he lowestpriority after transponder failure indications. When anew transponder message is received by the controlpanel, the “GO” annunciator illuminates.

2 - ALPHA-NUMERIC CHARACTER LCD

The Mode-S/IFF control panel has an eight alphanumeric character,liquid crystal display. The display has white characters on a blackbackground.

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3 - GO ANNUNCIATOR

The Mode-S/IFF control panel has a dead front green light GOannunciator to indicate no failures in the system or to indicate a newtransponder message has been received by the control panel.

4 - NO-GO ANNUNCIATOR

The Mode-S/IFF control panel has a dead front amber light NO-GOwarning annunciator to indicate system failure.

5 - TEST BUTTON

The TCAS/IFF control panel is provided with a momentary push-buttontest to activate the System Functional Test.

The Mode-S/IFF Transponder System test may be accomplished withthe Mode/TCAS rotary knob in the TA or TA/RA position and pressingthe TEST push button. In addition, the TCAS is tested when the test isperformed.

Pressing the TEST push button the Mode-S/IFF control panel starts alamp test, turning all front panel’s LCD indicators segments on.Additionally, a control panel self-test is performed when the TEST pushbutton is pressed for more than five seconds.

6 - MASTER CONTROL KNOB

The Master Control knob has OFF, STBY, NORM, and EMERpositions. A mechanical interlock prevents inadvertent switching toeither EMER or OFF positions.

OFF: In the OFF position all primary power supplies areremoved from the system, except control andlighting circuits.

STANDBY: In the STBY position the system disables thetransponder reply transmit (Modes 1, 2, 3/A, C, andS). STBY is typically engaged on the ground toprevent unnecessary RF traffic. It is disengagedjust prior to takeoff and engaged again uponlanding.


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NORMAL: The NORM position enables the transponder torespond to all proper modes (1, 2, 3/A, C, and S),as well as TCAS TA and TA/RA interrogations, ifselected.

EMERGENCY: The EMER position enables transponder modes 1,2, 3/A, and S to reply with an emergency codeautomatically, irrespective of settings of theMODE/TCAS rotary knob. Modes 1, 2, and 3/Aautomatically respond with a military emergency.Mode 3A replies is changed to code 7700.

7 - MODE/TCAS ROTARY KNOB

The MODE/TCAS rotary knob has six positions (OFF, 3A, C, S, TA,and TA/RA) that enables the different modes of operation and allowsfor selection of TCAS functions.

OFF: Disables the Modes 3A, C and S.

MODE 3/A: When Mode 3/A is enabled, replies to Mode 3/Ainterrogations and Mode C framing pulses areenabled.Mode 3/A control is provided for pilot or operatorselection of 4096 response codes. Code selectionrange is 0000 to 7777, inclusive. When the EMERmode is enabled, the Mode 3/A military emergencycode 7700 shall be displayed. The Mode 3/A coderemains at 7700 as long as the Master Control knobis in EMER position.

MODE C: When Mode C is enabled, the system replies to validMode C interrogations. The mode C control disablesor enables the Air Data Source input to thetransponder.

MODE S: This control function enables/disables Mode Soperation. When mode S is enabled, the systemreplies to valid ATCRBS/Mode S All-call and mode Sinterrogations.

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TCAS TA: This TCAS control function enables TA operationonly. The TA ONLY Mode-S enables the TCAScomputer, in conjunction with the Mode S/IFFtransponder, to provide traffic advisories. The TAMode is electrically tied to the Mode S enable (ON)mode. The system automatically selects this modewhen the aircraft is flying under 1000 ft aboveground level.

TCAS TA/RA: The TCAS TA/RA Mode S enables the TCAScomputer, in conjunction with the mode-S/IFFtransponder, to provide traffic and resolutionadvisories. TA/RA is electrically tied to the Mode Senable (ON) mode. This is the normal operationmode, supplying full TCAS coverage.

8 - CODE ENTRY KEYPAD

There are 12 push-button keys in the keypad.

9 - TCAS HORIZONTAL RANGE ROTARY KNOB

The TCAS range control allows the pilot to select traffic advisoryhorizontal display range in nautical miles. This control function givesthe pilot the option of selecting four possible ranges: 6, 12, 20, 40 NM.

10 - TCAS VERTICAL RANGE ROTARY KNOB

This knob provides the operator with the capability to select an altituderange, in relation to the airplane, for traffic display on the MFD.

ABV Mode: Range limits are 9900 ft above and 2700 ft belowairplane.

BLW Mode: Range limits are 9900 ft below and 2700 ft aboveairplane.

NORM Mode: The display range is 2700 ft above and below theairplane.

NOTE: If the TCAS does not receive the altitude limit information fromthe transponder, the default altitude limit presented will be7000 ft instead of 9900 ft.


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11 - IDENT MIC/OFF SWITCH

The Mode-S/IFF control panel has a two-position locking lever toggleswitch. The control provides the means to activate Ident functionremotely by keying the operator’s microphone (MIC position) orpressing the front panel IDNT push-button switch (OFF position).

12 - IDNT PUSH-BUTTON

A control panel function is supplied for enabling Identification ofPosition (I/P) operation, either directly by the operator or remotely bykeying the operator’s microphone. Momentarily pressing the IDNTbutton control causes the ATCRBS reply and/or Mode S Uplink (UF)-4and UF-5 replies to contain a Special Position Identifier (SPI) as anidentifier to the ground. The SPI supplies the ground station with morepositive identification of airplane position. This is typically activatedupon verbal command of the Air Traffic Controller. The IDNT control isa momentarily activated button to inhibit continuous selection. Whenpressure is removed from the momentary button, the IDNT controlreturns to the OFF position.

13 - LOAD PUSH-BUTTON

The LOAD push-button is used to transmit a completed code displayon the LCD.

14 - MODE 1/2 ROTARY KNOB

OFF: Disables Modes 1 and 2.

MODE 1: Enables Mode 1 operation. Mode 1 control is providedfor pilot or operator selection of 32 response codes.Code selection range is 00 to 73, inclusive.

MODE 2: Enables Mode 2 operation. Mode 2 control is providedfor pilot or operator selection of 4096 response codes.Code selection range is 0000 to 7777, inclusive.

MODE 1+2: Enables Modes 1 and 2 operation.

15 - MODE 4 CONTROL SWITCHES AND KNOB

The Mode 4 operation controls are disabled for this airplane.

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PRECISION AREA NAVIGATION (P-RNAV)P-RNAV defines European RNAV operations, which satisfy a requiredtrack-keeping accuracy of ±1 NM for at least 95% of the flight time,path coding in accordance with ARINC 424 (or an equivalentstandard), and the automatic selection, verification and, whereappropriate, de-selection of navigation aids.P-RNAV operations determines aircraft position in the horizontal planeusing inputs from the following types of positioning sensors (in nospecific order of priority):

− Distance Measurement Equipment (DME) giving measurementsfrom two or more ground station (DME/DME);

− VHF Omni-directional Range (VOR) with a co-located DME(VOR/DME), where it is identified as meeting the requirements of theprocedures;

− Global Navigation Satellite System (GNSS);− Inertial Navigation System (INS) or Inertial Reference System (IRS),

with automatic updating from suitable radio based navigationequipment.

P-RNAV is used for departures, arrivals and approach (FAWP - FinalApproach Waypoint), and not used on final approach, i.e. from FAWPto RWY and missed approach.

LIMITATIONS

− For P-RNAV operations in terminal airspace, obstacle clearanceprotection, up to the FAWP, will assume that aircraft comply with theP-RNAV accuracy requirements;

− Obstacle clearance altitude has been based upon the infrastructuregiving the poorest precision;

− The minimum flight crew are 2 Pilots;− It is not permissible to use, for any period of time, data from an

inertial system as the only means of positioning;− The system must display essential information in the Pilot’s primary

field of view such as:− Lateral Deviation;− TO/FROM waypoints;− Failure flag (failure of P-RNAV system);


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− Unless automatic updating of the actual departure point is provided,the flight crew must ensure initialization on the runway either meansof a manual runway threshold or intersection update, as applicable.This is to preclude any inappropriate or inadvertent position shift aftertake-off;

− Where reliance is placed on the use of radar to assist contingencyprocedures, its performance has been shown to be adequate for thatpurpose, and the requirement for a radar service is identified in theAIP;

− P-RNAV operations must use FMS to control all lateral navigationfunctions. For FMS limitations, refer to Limitations Section 1-01-60(System: FMS) of AOM;

− The system must have means to display to the flight crew thefollowing items:

− The active (TO) waypoint and distance/bearing to this point;− Ground speed or time to the active (TO) waypoint;− Automatic tuning of VOR and DME navigation aids used for

position updating together with the capability to inhibitindividual navigation aids;

− RNAV system failure;− Alternate means of displaying navigation information, sufficient

to perform cross-checks procedures.

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NORMAL PROCEDURES

− Verify NOTAM (Notice to Airman) for non-available P-RNAVprocedure, if navigational aids, identified in the AIP as critical fora specific P-RNAV procedure, are not available;

− Use phraseology appropriate to P-RNAV operations;− When the VOR or DME is not available or shutdown, the flight

crew have to inhibit the navigation aid from the automaticselection process;

− The flight crew must notify ATC of any problem with the RNAVsystem that results in loss of the required navigation capability,together with the proposed course of action;

− Discrepancies that invalidate a procedure must be reported tothe navigation database supplier and affected procedures mustbe prohibited by an operator´s notice to its flight crew.

PRE-FLIGHT PLANNING

− Verify the required navigation aids critical to the operation ofspecific procedure, and if they are identified in the AIP(Aeronautical Information Publication) and on the relevantcharts;

− Check availability of the navigation infrastructure and onboardequipment for the period of intended operation;

− The navigation database must be appropriate for the region ofthe intended operation and must include the navigation aids,waypoints, and coded terminal airspace procedures for thedeparture, arrival and alternate airfields;

− When specified in the AIP that dual P-RNAV procedure arerequired for specific terminal P-RNAV procedure, the availabilityof dual P-RNAV system must be confirmed;

− If a stand-alone GPS is to be used for P-RNAV, the availabilityof RAIM must be confirmed;

DEPARTURE

− Both Pilots must verify if the navigation database is current andif aircraft position has been entered correctly;

− The PNF (Pilot Not Flying) must verify the desired path and theaircraft position relative to the path;

− The active flight plan should be checked by comparing thecharts with the MAP display and the MCDU;

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− A procedure shall not be used if doubt exists as to the validity ofthe procedure in the navigation database;

− The creation of new waypoints by manual entry into the RNAVsystem by the flight crew is not permitted;

− Route modifications in the terminal area may take form of radarheadings or direct to clearances;

− Prior to take off, the flight crew must verify that the R-NAVsystem is available and operating correctly and, whereapplicable, the correct airport and runway data have beenloaded;

− Unless automatic updating of the actual departure point isprovided, the flight crew must ensure initialization on the runwayeither by means of a manual runway threshold or intersectionupdate, as applicable. This is to preclude any inappropriate orinadvertent position shift after take-off. Where GNSS is used,the signal must be acquired before the take-off roll commencesand GNSS position may be used in place of the runway update;

− During the procedure and where feasible, flight progress shouldbe monitored for navigational reasonableness, by cross-checks,with conventional aids using the primary displays in conjunctionwith the MCDU;

− When automatic update for departure is not available, theprocedure should be flown by conventional navigation means. Atransition to the P-RNAV structure should be made at the pointwhere the aircraft has entered DME/DME coverage and has hadsufficient time to achieve an adequate input. If a procedure isdesigned to be started conventionally, then the latest point oftransition to the P-RNAV structure will be marked on the charts.If a Pilot elects to start a P-RNAV procedure using conventionalmethods, there will not be any indication on the charts of thetransition point to the P-RNAV structure.

ARRIVAL

Prior to the arrival phase, the flight crew should verify that the correctterminal procedure has been loaded. The active flight plan should bechecked by comparing the charts with the MAP display and the MCDU.This includes confirmation of the waypoint sequence, reasonablenessof track angles and distances, any altitude or speed constraints, and,where possible, which waypoints are fly-by and which are fly-over.


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− If required by procedure, a check will need to be made toconfirm that updating will exclude a particular navigation aid. Aprocedure shall not be used if doubt exists as to the procedurein the navigation database;

− Where the contingency to revert to a conventional arrivalprocedure the flight crew must make the necessary preparation;

− During the procedure and where feasible, flight progress shouldbe monitored for navigational reasonableness by cross-checkswith conventional navigation aids using the primary displays inconjunction with the MCDU. In particular, for a VOR/DME RNAVprocedure, the reference VOR/DME used for the construction ofthe procedure must be displayed and checked by the flight crew.For RNAV systems without GNSS updating, a navigationreasonableness check is required during the descent phasebefore reaching the Initial Approach Waypoint (IAWP). ForGNSS based systems, absence of an integrity alarm isconsidered sufficient. If the check fails, a conventionalprocedure must then be flown;

− Route modifications in the terminal area may take the form ofradar headings or direct to clearances and the flight crew mustbe capable of reacting in a timely fashion. This may include theinsertion of tactical waypoints loaded from the database. Manualentry or modification by the flight crew of a loaded procedure,using temporary waypoints or fixes not provided in the database, is not permitted;

− Although a particular method is not mandated, any publishedaltitude or speed constraints must be observed.

CONTINGENCY PROCEDURES

− The flight crew must notify ATC of any problem with the RNAVsystem that results in the loss of required navigation capability,together with the proposed course of action;

− In the event of communication failure, the crew should continuewith the RNAV procedure in accordance with the published lostcommunication procedure;

− In case of loss of P-RNAV capability, the flight crew shouldnavigate using an alternative means of navigation. The alternatemeans need not be an RNAV system;

− Cautions and warnings for the following conditions:− Failure of the RNAV system components including those

affecting flight technical error;

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− Flight director – discontinue the P-RNAV procedure followingthe approved missed approach procedure or if feasible revertto a conventional or IRS procedure and inform ATC;

− Automatic Flight – continue the approach using manual flight,and if the flight path cannot be followed perform a approvedmissed approach procedure and inform ATC;

− Multiple system failures – If a multiple system failures occurssuch as affecting GNSS, Flight Director, and any other used forP-RNAV procedure, a missed approach procedure must beperformed and inform ATC;

− Failure of navigation sensors - discontinue the P-RNAVprocedure following the approved missed approach procedureor if feasible revert to a conventional or IRS procedure andinform ATC.

INCIDENT REPORTING

Significant incidents associated with the operation of the aircraft whichaffect or could affect the safety of RNAV operations, need to bereported on the appropriate report manifest.Specific examples may include:

− Aircraft system malfunctions during P-RNAV operations whichlead to:− Navigations errors not associated with transitions between

different navigation modes;− Significant navigation errors attributed to incorrect data or a

navigation database coding error;− Unexpected deviations in lateral or vertical flight path not

cause by Pilot input;− Significant misleading information without a failure warning;− Total loss or multiple navigation equipment failure;

− Problems with ground navigational facilities leading to significantnavigational errors not associated with transitions betweendifferent navigation modes.


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PNEUMATICSAIR CONDITIONING

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SECTION 2-14

PNEUMATICS, AIR CONDITIONINGAND PRESSURIZATION

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-14-05 ..01Pneumatic System ............................................................. 2-14-05 ..02

EICAS Messages ........................................................... 2-14-05 ..05Air Conditioning System..................................................... 2-14-10 ..01

ECU Operation ............................................................... 2-14-10 ..02Cabin Temperature Control............................................ 2-14-10 ..04Air Conditioning Distribution ........................................... 2-14-10 ..06Pneumatic System Function Logic ................................. 2-14-10 ..07EICAS Messages ........................................................... 2-14-10 ..10Controls and Indicators................................................... 2-14-10 ..11

Pressurization System ....................................................... 2-14-15 ..01Operation in Automatic Mode......................................... 2-14-15 ..02Operation in Manual Mode ............................................. 2-14-15 ..11EICAS Messages ........................................................... 2-14-15 ..11Controls and Indicators................................................... 2-14-15 ..12Pressurization Indication on EICAS................................ 2-14-15 ..14

Electronic Bay Cooling System .......................................... 2-14-20 ..01Forward Electronic Bay .................................................. 2-14-20 ..01Rear Electronic Bay........................................................ 2-14-20 ..02EICAS Messages ........................................................... 2-14-20 ..02

Baggage Ventilation System .............................................. 2-14-25 ..01

PNEUMATICSAIR CONDITIONINGAND PRESSURIZATION

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GENERALThe pneumatic system can be supplied by the engines, APU or aground pneumatic source.

The APU or ground pneumatic source supplies the system prior to theengine start. The engines normally supply bleed air for pneumaticsafter engine start.

The air conditioning system receives air from the pneumatic systemand provides conditioned air to the cabin. The system is controlled bytwo Environmental Control Units (ECU).

The pressurization system uses bleed air from the air conditioningsystem to pressurize the airplane. Cabin pressure is controlled bymodulating the outflow valves. The system is controlled by anautomatic mode and has a manual back-up mode.

Cooling for rear and forward electronic compartments is provided bythe ventilation system.

System information and messages are presented on the EICAS.


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PNEUMATIC SYSTEMThe pneumatic system receives compressed and hot air from thefollowing sources:− Engines compression stage;− APU;− Ground pneumatic source.

The pneumatic system is used for: engine start, air conditioning,pressurization and anti-ice system.

Engine bleed air comes from the 9th (low pressure) and 14th (highpressure) engine stages.

The 14th stage High Stage Valve (HSV), which is electrically controlledand pneumatically-actuated, opens automatically during low enginethrust operations, engine cross bleed start and anti-ice operation.

As thrust increases, the HSV closes and the 9th BACV (Bleed AirCheck Valve) opens supplying bleed air to the system.

Bleed air for engine anti-ice system is provided through the tappingupstream of the HSV.

An Engine Bleed Valve (EBV), which is electrically controlled throughthe Bleed Air Button and pneumatically-actuated, is installeddownstream of the pre-cooler.

Bleed air for the Air Turbine Starter (refer to Section 2-10 - Powerplant)is provided through the tapping downstream of the EBV.

Each engine supplies air to its corresponding air conditioning pack andanti-ice system.

A Cross-Bleed Valve (CBV), which is electrically controlled through theCross Bleed Knob and pneumatically actuated, provides thesegregation or interconnection between both sides in case of APUoperation or one engine pneumatic supply.

The pneumatic system’s functional logic opens or closes automaticallythe EBV, if the Cross Bleed Knob is on AUTO position, during enginestart, in order to select the available pneumatic source: APU, groundpneumatic source or opposite engine.

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The functional logic also opens automatically the CBV and both HSVand closes one air conditioning pack whenever the anti-icing system isoperating.

Bleed air from the APU, that is used primarily as a auxiliary pneumaticsource, is provided in the left side of the pneumatic system to supplythe air conditioning and engine starting either on ground or inflight.

An APU Bleed Valve (ABV), which is electrically controlled through theAPU Bleed Button and pneumatically-actuated, provides APU bleedcontrol.

The pneumatic system functional logic automatically closes the ABVwhenever any engine is supplying bleed air to the left pneumatic side.

An APU Check Valve is installed downstream of the APU bleed valve.

A ground pneumatic source connection, including a check valve, isinstalled on the right side of the pneumatic system. Its main purpose isto supply pressurized air to start the engines.

Leak detectors (thermal switches) are installed along all the pneumaticlines. Should a duct leakage occur, these detectors activates a warningmessage in the EICAS.

Should any hot air leakage occur, the bleed sensors and/or threeMassive Leakage Detectors (thermal switches – the former locatedalong the pneumatic system ducting and the latter in the rear electroniccompartment area) will close the EBV and the HSCV of the affectedside, as well as the CBV.

Bleed temperatures upstream and downstream of the pre-cooler aremonitored through temperature sensors. Temperature downstream ofthe pre-cooler is presented on a vertical bar indication on the MFD.

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EICAS MESSAGES


WARNING

BLD 1 (2) LEAKBLD APU LEAK

Duct leakage in theassociated bleed line.Temperature in the ductregion exceeds 91°C (195°F).The switch deactivates at79°C (175°F).

BLD 1 (2) OVTEMP Associated pre-cooler downs-tream temperature above305°C (581°F).

APU BLD VLV FAIL Disagreement between actualposition and commandedposition of the APU BleedValve.

BLD 1 (2) LOW TEMP Abnormal low or asymmetricbleed temperature, or pre-cooler outlet temperaturesensor failure.

CAUTION

BLD 1 (2) VLV FAIL Disagreement between actualposition and commandedposition of the associatedEngine Bleed Valve.

CROSS BLD FAIL Disagreement between actualposition and commandedposition of the Cross-BleedValve.

CROSS BLD SW OFF Cross Bleed Knob selectedCLOSED with at least oneengine running after brakerelease.

HS VLV 1 (2) FAIL Disagreement between actualposition and commandedposition of the associatedHigh Stage Valve.

ADVISORY

BLD 1 (2) VLV CLSD Associated Engine BleedValve position. This messageis inhibited on ground orduring associated enginestart.

CROSS BLD OPEN Cross Bleed Valve open.


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AIR CONDITIONING SYSTEMAirplane air conditioning is provided by two the Environmental ControlUnits (ECU) supplied by the Pneumatic System.

Each side is provided with independent controls, protection devices,and cross-connected air distribution lines for the various modes ofoperation.

Cockpit and passenger cabin temperature selections are independentand may be controlled either manually or automatically. The left ECUcontrols the temperature in the cockpit and the right ECU controls thetemperature passenger cabin.

The system is normally operated in the automatic mode. In case ofautomatic mode failure, a manual mode is available.

The pilots may transfer the passenger cabin temperature control to theAttendant Panel.

The air conditioning distribution is performed by the gasper system andgeneral outlets with cross-connection between the cockpit andpassenger cabin lines.

This feature, associated with the ram air inlets, allows the cockpit andpassenger cabin to be supplied with fresh air, in case of failure of bothECUs.

Recirculating air, driven by two electrical fans, is mixed to fresh air inorder to improve passenger and crewmembers' comfort.

A ground cart connection is available at the right-hand duct, connectedto the outside through a check valve in the fuselage. The pre-conditioned air from the ground cart is delivered to the cabin directlythrough the distribution lines.

The air conditioning system incorporates protection features in thetemperature controllers which shut off the system in case ofmalfunctions (duct leakage, duct overtemperature, and packovertemperature).

The cockpit and passenger cabin temperature indications arepresented on the MFD. Caution and advisory messages are presentedon the EICAS.


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ECU OPERATION

Each ECU consists of a dual heat exchanger, an air cycle machine(compressor, turbine, and fan), a condenser, a water separator andrelated control and protective devices, installed forward of the airplanewing root, inside the wing-to-fuselage fairing.

The automatically-controlled bleed air from the pneumatic systemsupplies the ECU. Downstream pressure is regulated by the PackValve (Pressure Regulating and Shutoff Valve).

After the Pack Valve, the airflow is divided into two lines:

- One cold line that passes through to the Air Cycle Machine.

- One hot line that bypasses the Air Cycle Machine.

Both airflow lines are gathered at the expansion turbine discharge.

In the Air Cycle Machine (ACM), air is cooled in the primary heatexchanger and passes through the compressor, thus causing apressure increase. The air then goes to the secondary heat exchangerwhere it is cooled again.

After leaving the secondary heat exchanger, the high-pressure cooledair passes through a condenser and a water separator for condensedwater removal. Spray nozzles uses the separated water to improve theheat exchanger efficiency.

The main airstream is ducted to the turbine and expanded to providepower for the compressor and cooling fan. This energy removalproduces very low turbine discharge temperatures, achieving adequatelow temperatures in the process.

The cold exit air is mixed with warm air supplied by the recirculationfan and/or with the hot bypass air immediately upon leaving the turbine.A check valve is provided in the recirculation duct to prevent reverseflow if the recirculation fan is inoperative.

The ECU outlet air temperature is controlled through the dualtemperature control valve. One valve adds hot bleed air to the turbinedischarge while the other valve restricts the compressor inlet flow.

The ECUs are cooled in flight by external the ACM fans, using theexternal ram air. On the ground, the ECUs are cooled by the ACM fansonly.

The system has emergency ventilation, as an alternate means to allowthe outside air into the cabin. The impact air passes through the sameram air inlets that are used to cool the dual heat exchangers.

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When the ECUs air supply is shut off in flight, the emergency ram air isactivated and the ram air valves are opened automatically, allowingram air to be routed to the distribution lines. Ram air may also be usedto ventilate the airplane interior for cabin smoke evacuation and cabinventilation purposes with the airplane depressurized and the ECUsturned off.

NOTE: The Pneumatic System automatic logic closes the left PackValve whenever the anti-icing system is operating below24600 ft.

CABIN TEMPERATURE CONTROL

AUTO MODE

In the automatic mode (temperature knobs pressed), the temperaturesin the passenger cabin and in the cockpit are controlled by the digitaltemperature controllers that receive information from the temperaturesensors (ducts, passenger cabin, or cockpit), maintaining thetemperature set on the associated temperature knob.

MANUAL MODE

In manual mode (temperature knobs pulled), the temperature in thepassenger cabin and in the cockpit are controlled by the temperaturecontrol module, that receives information from the temperature knobsand the duct temperature sensor.

The manual mode should be used only if a failure occurs in theautomatic mode and may be noticed when the temperature is notmaintained within the temperature limits of the automatic mode(between 18 and 29°C) after cabin temperature stabilization.If switching from auto mode to manual mode is required, proceed asfollows:• Set the knob to mid range position (12 o’clock).• Wait for system to stabilize (approximately 30 seconds).• Switch to manual.• Smoothly turn the knob to the required point.

Once in the manual mode, the pilot must continuously monitor thetemperature and actuate on the Temperature and Mode Selector Knob.

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AIR CONDITIONING DISTRIBUTION

The air conditioning distribution system provides conditioned air to thecockpit and passenger cabin.

The main source of conditioned air to the cockpit is the left pack, with asingle distribution system for cooling or heating air.

The cockpit is provided with two FEET AIR handles and air outlets,allowing each pilot to individually control the airflow.

For CRT displays ventilation, a shutoff valve on each side, electrically-driven and independently controlled by a thermal switch, allows cold airto be supplied for this function only.

The main source of conditioned air to the passenger cabin is the rightpack and partially by the left pack, through a cross connection duct.The air distribution system for the passenger cabin is divided into threelines. One line is distributed to the lower ducts, installed at the footlevel on both cabin sidewalls. The second line is for the upper ducts ofboth sidewalls. The third line is dedicated to the gasper. If the ducttemperature is below 24°C (75°F), the associated temperatureswitches command the recirculation fans to increase airflow.

The gasper air subsystem provides air to individual air outlets (gasper),as well as for the rear electronic compartment, oxygen cylindercompartment and relay box ventilation. The air to the gasper isprovided by a gasper fan and by one branch from the cross connectionof the general distribution system. The gasper fan is similar to therecirculation fan, but it is operated in normal condition only. Onethermal switch is installed in the branch line to close fresh air in case ofheating condition (above 24°C). In this case, only air from the gasperfan is available.

The recirculation air subsystem, consists of two recirculation fans, andis usually operated to save the engine bleed. It must be kept off shouldthere be smoke in the cabin, or on hot days while on the ground. Thisreduces the pull-down period and should be turned on in cold soakconditions to reduce pull-up period.

The operational logic to open the Engine Bleed, Cross-bleed, APUBleed, and Pack Valves will be analyzed herein separately, for bettersystem comprehension. This system also actuates on the Anti-icingSystem Valves. For further information, refer to Section 2-15 - Ice andRain Protection.

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PNEUMATIC SYSTEM FUNCTIONAL LOGIC

The pneumatic system functional logic provides automatic control andprotection for itself and the user systems, giving priority according tothe airplane operation or condition.

ENGINE BLEED VALVE LOGIC

The Engine Bleed Valve (EBV) receives an electrical input to openwhen the following conditions occur simultaneously:

− Bleed Air Button is pressed to open the valve;− Respective Essential Bus is energized;− There is no massive leakage on the respective side of the rear

electronic compartment;− There is no leakage along the pneumatic system ducting;− Respective engine N2 is above 56.4%; and− Respective engine fire extinguishing handle is not pulled. APU BLEED VALVE OPERATIONAL LOGIC

The APU Bleed Valve (ABV) receives an electrical input to open whenthe following conditions occur simultaneously:

− APU Bleed Button is pressed to open the valve;− Essential DC Bus 1 is energized;− Engine 1 bleed valve is closed (no pressure from the left side);− Engine 2 bleed valve or cross-bleed valve is closed (no pressure

from the right side);− APU rpm above 95%, plus 7 seconds; and− There is no massive leakage on the APU line.


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PACK VALVE OPERATIONAL LOGIC

The Pack Valve receives an electrical input to open when the followingconditions occur simultaneously:

− Air Conditioning Pack Button is pressed to open the valve;− Respective DC Bus is energized;− Respective engine is not starting;− No engine is starting using the APU as pneumatic source;− No failure in the related pack is detected (overpressure,

overtemperature or duct leakage downstream of the Pack Valve);and

− No discrete ECS (Environmental Control System) OFF signal issent from any related FADEC (A or B).

The FADEC`s discrete ECS OFF signals are produced according tothe following conditions:

1- During Takeoff or Go Around:

ACTIVATION CONDITIONS FOR ECS OFF SIGNALS

PRESSURE ALTITUDE / TAT °C

ENGINE FADEC MODEALL ENGINESOPERATIVE(takeoff only)

ONEENGINE

INOPERATIVE (3)

A1P ALLALT T/O-1T/O-1 orT/O RSV

Up to 1700 ft abovetakeoff altitude (1)

Lower than9700 ft (2)

A1E ALL

ALT T/O-1T/O-1,E T/O,

T/O RSV orE T/O RSV

Up to 1700 ft abovetakeoff altitude (1)

Lower than9700 ft (2)

NOTE: 1) TAT above 19°C (66°F) at sea level, decreasing linearly to−5°C (23°F) at 9700 ft.

2) The ECS OFF signal is activated for the Pack associatedwith the operating engine if the pressure altitude is lowerthan 9700 ft and TAT is above 19°C at sea level,decreasing linearly to −5°C at 9700 ft (area A in thefollowing envelope).

3) A Low N1 condition (actual N1 does not achieve requestedN1) is considered one engine inoperative.

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FADEC´S ECS OFF ENVELOPE

The ECS OFF logic is valid only when the packs are using enginebleed. If APU bleed is being used, the ECS OFF logic is inhibited andthe pack valves will not shut down.

The FADEC´s discrete ECS OFF signal is not produced when usingALT-T/O-1 mode during takeoffs with all engines operative.

If a FADEC commands its associated pack to close, through the ECSOFF signal, the pilot must reset the pack when the conditions for theautomatic shut down of the pack cease to exist, i.e., an automaticrestart of the pack does not exist.

For airplanes S/N 145462, 145516, 145591, 145644 and on, the packsare automatically reset when the conditions for the ECS OFF signalcease to exist. When both packs are automatically reset, pack 2 will becommanded to open 10 seconds after pack 1 opening, to avoidpassenger discomfort due to packs return.

2- During reverse use:

− The ECS OFF signal is always activated during reverse use.

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CROSS BLEED VALVE OPERATIONAL LOGIC

The Cross-Bleed Valve (CBV) receives an electrical input to openwhen the following conditions occur:

− Essential DC Bus 2 is energized;− There is no leakage along the pneumatic system ducting or a

massive leakage in the Rear Electronic Compartment; and− Cross-Bleed Knob is set to OPEN; or− Cross-Bleed Knob is set to AUTO and one of the following

conditions occurs:− Horizontal Stabilizer Anti-Icing System is operating; or− Engine 2 is starting; or− Engine 1 is starting assisted by engine 2 or external pneumatic

source (with APU Bleed Valve manually commanded to theclose position).

EICAS MESSAGES

TYPE MESSAGE MEANING PACK 1 (2) OVLD Associated ECU compressor

temperature above 243°C(470°F) or ECU inlet pressureabove 55 psig.

CAUTION PACK 1 (2) OVHT Associated ECU outlet

temperature above 93°C(200°F).

PACK 1 (2) VLV FAIL RAM AIR VLV FAIL

Disagreement betweenassociated valve actualposition and commandedposition.

ADVISORY

PACK 1 VLV CLSD Left pack valve closed with noicing condition, or Left pack valve closed withairplane above 24600 ft.

PACK 2 VLV CLSD Right pack valve closed.

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AIR CONDITIONING AND PNEUMATIC CONTROL PANEL

1 - COCKPIT TEMPERATURE AND MODE SELECTOR KNOB− PRESSED - Controls the left pack in automatic mode through

the Digital Temperature Controller. The cockpit temperaturemay be set between 18°C (65°F) and 29°C (85°F).

− PULLED - Controls the left pack in manual mode through thetemperature control module. No temperature range isestablished.

2 - PASSENGER CABIN TEMPERATURE AND MODE SELECTOR

KNOB− PRESSED - Controls the right pack in automatic mode through

the Digital Temperature Controller. The passenger cabintemperature may be set between 18°C (65°F) and 29°C (85°F).

− PULLED - Controls the right pack in manual mode through themanual mode circuit in the temperature control module. Notemperature range is established.

− ATTD - The passenger cabin temperature control is transferredto the attendant’s panel in automatic mode only.

3 - RECIRCULATION BUTTON

− Turns on (pressed) or turns off (released) both recirculationfans.


4 - AIR CONDITIONING PACK BUTTON

− Opens (pressed) or closes (released) the Pressure Regulatingand Shutoff Valve of the associated ECU.



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5 - GASPER BUTTON− Turns on (pressed) or turns off (released) the gasper fan inflight

only.− A striped bar illuminates inside the button to indicate that it is

released.− On ground, the gasper fan is turned on as soon as the

associated DC Bus is energized. 6 - CROSS-BLEED KNOB

− CLOSED- Closes the Cross-bleed Valve.− AUTO - Selects automatic operation mode of the Cross-bleed

Valve.− OPEN - Opens the Cross-bleed Valve.

7 - BLEED AIR BUTTON

− Opens (pressed) or closes (released) the associated EngineBleed Valve.


− A LEAK inscription illuminates inside the button to indicate aduct leakage in the associated bleed line.The LEAK inscription is not available on some airplanes.

8 - APU BLEED BUTTON− Opens (pressed) or closes (released) the APU Bleed Valve.− A striped bar illuminates inside the button to indicate that it is

pressed.− An OPEN inscription illuminates inside the button to indicate that

the APU Bleed Valve is in the open position.

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ENVIRONMENTAL CONTROL SYSTEM (ECS) ANDPNEUMATIC PAGE ON MFD

1 - PASSENGER CABIN TEMPERATURE INDICATION− Indicates the temperature inside the passenger cabin.− Digits are green.− Legends are white.− Ranges from –10 to 50°C (14 to 122°F).

2 - COCKPIT TEMPERATURE INDICATION− Indicates the temperature inside the cockpit.− Digits are green.− Legends are white.− Ranges from –10 to 50°C (14 to 122°F).

3 - BLEED TEMPERATURE INDICATION− Indicates the bleed air temperature downstream of the pre-

cooler on the left and right engine.− Scale and Pointer:

− White for the scale, below 260°C (500°F) to indicatepotentially low thermal energy availability to the anti-icingsystem. Amber for the pointer, only if the pointer is in thewhite band of the scale and the message “BLD 1 (2) LOWTEMP” is shown on EICAS.

If the pointer is in the white band of the scale and themessage “BLD 1 (2) LOW TEMP” is not presented in theEICAS, the pointer will be green.

− Green from 260 to 305°C (500 to 581°F) to indicate theacceptable range.

− Red above 305°C (581°F) to indicate an overtemperaturecondition.

− In case of an outlet temperature sensor failure, the respectivepointer is removed from the vertical temperature bar.

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01 JUNE 20, 2002

GALLEY CONTROL PANEL

1 - CABIN TEMPERATURE INDICATION− Indicates the cabin temperature in bargraph format.

2 - CABIN TEMPERATURE CONTROL

− Increase or decrease the cabin temperature through the s andt keys.

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PRESSURIZATION SYSTEM

The Cabin Pressure Control System (CPCS) controls the cabinpressure by regulating the cabin air exhaust rate supplied by theECUs.

The CPCS comprises two subsystems:

- One digital electropneumatic subsystem (automatic mode).

- One pneumatic subsystem (manual mode).

The Cabin Pressure Control System comprises a digital controller, amanual controller, an electropneumatic outflow valve, a pneumaticoutflow valve, an air filter, two pressure regulator valves, an ejectorpump, two static ports, and a Cabin Pressure Acquisition Module(CPAM).

Both outflow valves receive static pressure signals from static ports foroverpressure relief and negative pressure relief functions, actuatingpneumatic devices to inhibit airplane structural damage or injury incase of improper system operation.

The safety devices provide the following features:

Airplanes up to S/N 854 and Pre-Mod. SB 145LEG-00-0007:− Positive cabin differential pressure relief: 8.4 psi maximum.− Negative cabin differential pressure relief: - 0.3 psi.− Cabin altitude limitation (when in the auto mode): 15000 ft maximum.

Airplanes S/N 863 and on or Post-Mod. SB 145LEG-00-0007:− Positive cabin differential pressure relief: 8.6 psi maximum.− Negative cabin differential pressure relief: - 0.3 psi.− Cabin altitude limitation (when in the auto mode): 15000 ft maximum.

The system is normally operated in the automatic mode. The manualmode is used in case of automatic mode failure.

The cabin air filter is provided to prevent nicotine and dust to enter theoutflow valve chamber.

Indications of cabin altitude, cabin differential pressure, and cabinaltitude rate of change are presented on the EICAS.

A caution message is presented on the EICAS in case of automaticmode failure, requiring the crew to select the manual mode.


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The CPAM and CPCS have internal tolerances of ± 100 ft and ± 200 ft,respectively. Then, depending on these tolerances accumulation, thedisplayed cabin altitude may be increased up to 300 ft.

If, however, the cabin altitude indication continuously increases and thesystem is out of its normal range of operation, causing a cabindepressurization, the CPAM sends a signal to the aural warningsystem to alert the crew when cabin altitude is above 9900 ± 100 ft.

OPERATION IN AUTOMATIC MODE

The automatic mode maintains minimum cabin altitude according tothe airplane operating altitude, imposing minimum cabin altitude rate ofchange.

The automatic mode is controlled by the digital controller and requiresa landing altitude to be entered prior to takeoff. According to thelanding altitude, the measured cabin pressure, ADC inputs (airplanealtitude, altitude rate of change and barometric correction), air/groundposition, and thrust lever position, the digital controller determines theadequate opening of the electropneumatic outflow valve.

On airplanes up to S/N 854 and Pre-Mod. SB 145LEG-00-0007, thepneumatic outflow valve is slaved to the electropneumatic outflowvalve and both operate simultaneously, maintaining the same positionwhile in the automatic mode.

On airplanes S/N 863 and on or Post-Mod. SB 145LEG-00-0007,during the operation on the automatic mode only the electropneumaticoutflow valve is actuated, being the pneumatic outflow valve closed.

Different operation sequences are automatically initiated by the DigitalController following the received inputs.

The Digital Controller schedules a cabin altitude that is the value thatthe measured cabin altitude must be equal to.

Cabin altitude rate of change varies according to the different operationsequences.

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Proper operation of the pressurization system in the automatic moderequires that the following conditions be met:

− Automatic mode is selected on the Digital Controller (button notpressed and MAN inscription not illuminated). The pressurizationsystem is in the automatic mode when electrical power is firstapplied.

− Landing altitude is entered in the Digital Controller prior to thetakeoff. Should the landing altitude not be entered, the system willautomatically consider 8000 ft as the landing altitude.

− Manual Controller is set to DN position (full counterclockwise). If theManual Controller is out of the DN position, the pneumatic valvetends to open causing inappropriate automatic mode operation.

DETERMINATION OF THE THEORETICAL CABIN ALTITUDE

The theoretical cabin altitude is a function of the airplane operatingaltitude. It is calculated in such a way that the maximum cabindifferential pressure is reached at the lowest possible airplane altitudeconsidering a minimum cabin altitude rate of climb and a maximumairplane rate of climb.

The maximum cabin differential pressure is 8.1 psi for airplanes up toS/N 854 and Pre-Mod. SB 145LEG-00-0007 and 8.4 psi for airplanesS/N 863 and on or Post-Mod. SB 145LEG-00-0007.

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AUTOMATIC PREPRESSURIZATION SEQUENCE ON GROUND

This sequence is initiated and maintained as long as the airplane is onthe ground and the thrust lever is set to THRUST SET position orabove.

For airplanes up to S/N 854 and Pre-Mod. SB 145LEG-00-0007, itcauses the cabin altitude to descend toward an altitude equivalent to400 ft (0.2 psi) below the takeoff altitude.

For airplanes S/N 863 and on or Post-Mod. SB 145LEG-00-0007, itcauses the cabin altitude to descend toward an altitude equivalent to750 ft (0.4 psi) below the takeoff altitude.

The purpose of the automatic pre-pressurization is to avoid cabinbumps due to the irregular airflow on the fuselage during rotation andtakeoff and also to keep a controlled cabin altitude just after rotation,as the cabin altitude tends to follow the airplane altitude.

In the case of takeoff with air conditioning supply, the cabin altitude iscontrolled with an altitude rate of descent equal to –450 ft/min.

In the case of takeoff without air conditioning supply, the outflow valvesare closed, also avoiding cabin bump.

TAKEOFF SEQUENCE

This sequence is initiated after the airplane leaves the ground with thepurpose of avoiding reselecting the landing altitude, in case it isnecessary to return to the takeoff airport.

For airplanes up to S/N 854 and Pre-Mod. SB 145LEG-00-0007, itcauses the cabin altitude to continue descending towards the altitudeequivalent to 400 ft below the takeoff altitude. If an altitude of 400 ftbelow the takeoff altitude has already been reached during thepre-pressurization sequence, the cabin altitude does not change.

For airplanes S/N 863 and on or Post-Mod. SB 145LEG-00-0007, itcauses the cabin altitude to continue descending towards the altitudeequivalent to 750 ft below the takeoff altitude. If an altitude of 750 ftbelow the takeoff altitude has already been reached during thepre-pressurization sequence, the cabin altitude does not change.

The takeoff sequence lasts until the theoretical cabin altitude becomesgreater than the actual cabin altitude, or until 15 minutes have elapsedsince the sequence initiation, whichever occurs first.

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FLIGHT SEQUENCE

This sequence is initiated after the takeoff sequence is finished, toestablish a cabin altitude and a cabin altitude rate of change duringflight.

For airplanes up to S/N 854 and Pre-Mod. SB 145LEG-00-0007, theDigital Controller schedules a cabin altitude that is the greatest valuebetween the theoretical cabin altitude and the selected landing altitudeminus 300 ft.

For airplanes S/N 863 and on or Post-Mod. SB 145LEG-00-0007, theDigital Controller schedules a cabin altitude that is the greatest valuebetween the theoretical cabin altitude and the selected landing altitudeminus 650 ft.

The cabin altitude rate of change is controlled at different valuesdepending on the scheduled cabin altitude and the airplane verticalspeed, but is limited to –450 ft/min during descent and 700 ft/min whileclimbing.

Barometric correction, when required, is automatically provided by theAir Data Computer (ADC).


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AUTOMATIC MODE OPERATION SCHEMATIC APPLICABLE TOAIRPLANES UP TO S/N 854 AND PRE-MOD. SB 145LEG-00-0007

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AUTOMATIC MODE OPERATION SCHEMATIC APPLICABLE TO

AIRPLANES S/N 863 AND ON OR POST-MOD. SB 145LEG-00-0007


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AUTOMATIC INCREASED RATE OF DESCENT SEQUENCE

This sequence is initiated when the airplane descent rate is greaterthan 200 ft/min, in order to satisfy all airplane rapid descent cases. Thecabin altitude rate of change limits may be accordingly increased,depending on the remaining flight time which is calculated consideringthe airplane operating altitude, airplane vertical speed and the selectedlanding altitude.

Therefore, the cabin altitude rate of descent limit may be increased toa value between –450 ft/min and –500 ft/min. In case the selectedlanding altitude is higher than 8000 ft, the cabin altitude rate of descentlimit may be increased to a value between +700 ft/min and+1500 ft/min.

AUTOMATIC DEPRESSURIZATION SEQUENCE ON GROUND

This sequence is initiated when the airplane is on the ground and thethrust lever is in the IDLE position.

To avoid a cabin bump during the landing, it is necessary that theairplane land with the cabin being submitted to a small differentialpressure.

For that reason, the automatic mode always controls, for landing, acabin altitude equal to the selected landing altitude minus 300 ft (forairplanes up to S/N 854 and Pre-Mod. SB 145LEG-00-0007) or 650 ft(for airplanes S/N 863 and on or Post-Mod. SB 145LEG-00-0007). Thissequence cancels this differential pressure corresponding to 300 ft or650 ft, as well as reduces cabin bump when the air conditioning isturned off or the main door is open.

Cabin depressurization is controlled at a rate of climb equal to650 ft/min, up to the full opening of the outflow valves.

In automatic mode, the rapid cabin depressurization is commanded bythe Dump Button.

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OPERATION IN MANUAL MODE

Manual operation is accomplished through the manual controller whichactuates only the pneumatic outflow valve, while the electropneumaticoutflow valve is kept closed, by selecting MAN in the PressurizationMode Selector Button and rotating the Manual Controller until thedesired cabin rate of change is reached. The crew is responsible formonitoring cabin differential pressure within acceptable values.

In manual mode, the DUMP button is not effective and a rapid cabindepressurization is commanded by turning the manual controller to theUP position (clockwise stop). In this mode, the cabin altitude limitationat 15000 ft does not exist as it does in the automatic mode.

EICAS MESSAGE


CAUTION PRESN AUTO FAIL Automatic pressurization mode

failure.


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DIGITAL CONTROLLER

1 - LANDING ALTITUDE INDICATOR− Displays the selected landing altitude.− Displays a failure code if any failure is detected during power-up

and continuous monitoring tests . In this case, the selection ofthe landing altitude is disabled.

− Successful power-up test is displayed (all light segmentsilluminated) until a landing altitude is selected.

− Displays blanks when Dump button or Mode Selector Button ispressed.

2 - LANDING ALTITUDE SELECTOR SWITCH− Sets the landing altitude in the Landing Altitude Indicator.− Altitude changes in 100-ft steps. Holding the selector for more

than 5 seconds changes the altitude in a 1000 ft/sec rate.− Landing altitude setting from –1500 ft to +14000 ft.

3 - PRESSURIZATION MODE SELECTOR BUTTON (guarded)− Provides selection of either automatic mode (button released) or

manual mode (button pressed) of operation.− When pressed, the MAN inscription illuminates inside the

button.

NOTE: In case of electrical failure that leads to the complete turningoff of the automatic mode turning off, manual mode shouldbe selected by pressing the Pressurization Mode SelectorButton, but the MAN inscription will not be illuminated.

4 - PRESSURIZATION DUMP BUTTON (guarded)− Provides rapid cabin depressurization up to 14500 ft.− When pressed, an ON inscription illuminates inside the button.− This button is effective in the automatic mode only.

MANUAL CONTROLLER KNOB

− Selects cabin rate of change between –1500 ft/min (at DN position)and approximately + 2500 ft/min (at UP position), when in themanual operating mode.

− When operating in the AUTO mode, it must be set to the DNposition.

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PRESSURIZATION CONTROLS AND INDICATORS


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PRESSURIZATION INDICATION ON EICAS

AIRPLANES UP TO S/N 854 AND PRE-MOD. SB 145LEG-00-0007

1 - CABIN ALTITUDE INDICATION− Displays cabin altitudes, regardless of the operating mode.− Ranges from – 1500 to 39000 ft, with a resolution of 100 ft.− Green: from – 1500 to 8300 ft.− Amber: from 8400 to 9900 ft.− Red: from 10000 to 39000 ft.

2 - DIFFERENTIAL PRESSURE INDICATION− Displays the differential pressure between the cabin interior and

the outside, regardless of the operating mode.− Ranges from – 0.5 to 10.0 psi, with a resolution of 0.1 psi.− Green: from 0.0 to 8.2 psi.− Amber: from – 0.3 to – 0.1 psi and from 8.3 to 8.6 psi.− Red: from – 0.5 to – 0.4 psi and from 8.7 to 10.0 psi.

3 - CABIN RATE OF CHANGE INDICATION− Displays the cabin rate of change, regardless of the operating

mode.− Ranges from –2000 to 2000 ft/min, with a resolution of 50 ft/min.− Green full range.− For rates out of range the indication is replaced by amber dashes.


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AIRPLANES S/N 863 AND ON OR POST-MOD. SB 145LEG-00-0007

1 - CABIN ALTITUDE INDICATION− Displays cabin altitudes, regardless of the operating mode.− Ranges from – 1500 to 41000 ft, with a resolution of 100 ft.− Green: from – 1500 to 8300 ft.− Amber: from 8400 to 9900 ft.− Red: from 10000 to 41000 ft.

2 - DIFFERENTIAL PRESSURE INDICATION− Displays the differential pressure between the cabin interior and

the outside, regardless of the operating mode.− Ranges from – 0.5 to 10.0 psi, with a resolution of 0.1 psi.− Green: from 0.0 to 8.5 psi.− Amber: from – 0.3 to – 0.1 psi and from 8.6 to 8.9 psi.− Red: from – 0.5 to – 0.4 psi and from 9.0 to 10.0 psi.

3 - CABIN RATE OF CHANGE INDICATION− Displays the cabin rate of change, regardless of the operating

mode.− Ranges from –2000 to 2000 ft/min, with a resolution of 50 ft/min.− Green full range.− For rates out of range the indication is replaced by amber dashes.



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ELECTRONIC BAY COOLING SYSTEM

FORWARD ELECTRONIC BAYAn automatic cooling system is provided in the nose electronic bay,where most of the electronic equipment is installed. This systemmaintains the temperature inside the bay within the avionicsoperational limits.The system comprises two NACA air inlets, two shutoff valves, tworecirculation fans, two exhaust fans, two check valves, four controlthermostats, and two overtemperature thermostats.

The NACA air inlets are provided with water separators and drains todeter water ingestion by the air inlets into the compartment.

All the fans are powered by four dedicated Inverter Modules.When the airplane is energized, the inverter modules are turned on,supplying power to the recirculation fans.The electrical power supply to the recirculation fan 2, exhaust fan 1and shutoff valve 1 is completely segregated from the remainingcomponents, to prevent a total loss of the system in case of anelectrical system single failure. Each recirculation fan operatescontinuously when its associated bar is energized.

A check valve is installed on each exhaust duct (left and right) to avoidwater ingestion through the exhaust fans.

If the forward electronic bay internal temperature exceeds 24°C (75°F)the control thermostats open the shutoff valves and turn the exhaustfans on. When the temperature drops below 19°C (66°F), the shutoffvalves are closed and the exhaust fans are turned off.

In the event that the temperature limit is reached, two overtemperaturethermostats are actuated and a caution message is presented on theEICAS.


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REAR ELECTRONIC BAYIn flight or during operation with the doors closed, rear electronic baycooling is performed by conditioned air discharged from the cabin.When this air flows from the underfloor area to the outflow valves,installed on the rear pressure bulkhead, it passes through thiscompartment, cooling it.

During ground operation, with the airplane unpressurized, an air outletblows air from the gasper fan line towards the rear electronic bay.

EICAS MESSAGE


CAUTIONELEKBAY OVTEMP Temperature inside the forward bay

exceeds 71ºC (160°F) maximum.

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FORWARD ELECTRONIC BAY COOLING SCHEMATIC


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BAGGAGE VENTILATION SYSTEM

Airplanes equipped with “class-C” baggage compartment have aBaggage Ventilation System installed. Although no dedicatedtemperature control is available (the “class-C” baggage compartmentis heated by the passenger cabin air flowing into it), the BaggageVentilation System provides an adequate environment for carrying liveanimals in the compartment.

The Baggage Ventilation System is composed of two ambient checkvalves and a baggage compartment fan.

Whenever the recirculation fan is off, the forward check valve preventsreverse flow into the baggage compartment and the two check valvesprevent smoke or fire extinguishing agent penetration into thepassenger cabin or into the rear electronic compartment, (refer toSection 2-7 - Fire Protection).


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SECTION 2-19

AUTOPILOT

TABLE OF CONTENTS

Block Page

General ............................................................................2-19-05.....01Automatic Flight Control System......................................2-19-05.....02Flight Guidance System ...................................................2-19-05.....04

Flight Director ...............................................................2-19-05.....04Autopilot........................................................................2-19-05.....04

Flight Director Modes .......................................................2-19-10.....01Lateral Modes...............................................................2-19-10.....01

Heading Hold Mode ..................................................2-19-10.....01Heading Select Mode (HDG) ....................................2-19-10.....02VOR NAV Mode (VOR) ............................................2-19-10.....03VOR Approach Mode (VAPP)...................................2-19-10.....04Localizer Mode (LOC/BC).........................................2-19-10.....04LNAV Mode ..............................................................2-19-10.....05

Vertical Modes..............................................................2-19-10.....06Pitch Hold Mode........................................................2-19-10.....06Altitude Hold Mode (ALT) .........................................2-19-10.....06Altitude Preselect Mode (ASEL) ...............................2-19-10.....07Flight Level Change Mode (FLC)..............................2-19-10.....07Speed Hold Mode (SPD) ..........................................2-19-10.....10Vertical Speed Hold Mode (VS) ................................2-19-10.....11Glide Slope Mode (GS).............................................2-19-10.....12Go Around Mode ......................................................2-19-10.....13Windshear Escape Guidance Mode .........................2-19-10.....15

Autopilot Disengagement .................................................2-19-10.....16EICAS Messages .............................................................2-19-15.....01Controls and Indicators ....................................................2-19-15.....01

Flight Guidance Controller............................................2-19-15.....01Pitch and Turn Controller..............................................2-19-15.....04Control Wheel...............................................................2-19-15.....05Thrust Levers ...............................................................2-19-15.....07Display Controller .........................................................2-19-15.....08PFD Indicators..............................................................2-19-15.....10EICAS Indicators ..........................................................2-19-15.....16

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Category II Approach....................................................... 2-19-20 .... 01Category II Conditions ................................................. 2-19-20 .... 01

Localizer Excessive Deviation Warning................... 2-19-20 .... 02Glideslope Excessive Deviation Warning ................ 2-19-20 .... 02

Controls and Indicators ................................................... 2-19-20 .... 02PFD Indicators ............................................................. 2-19-20 .... 02

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GENERAL

The Primus 1000 (P-1000) Automatic Flight Control System (AFCS) isa fully integrated, fail passive three-axis flight control system whichincorporates lateral and vertical guidance, yaw damper and automaticpitch trim functions.


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AUTOMATIC FLIGHT CONTROL SYSTEM

The Automatic Flight Control System (AFCS) consists of dual IC-600s,autopilot servos, a flight guidance controller (GC-550), a pitch and turncontroller (PC-400) and a display controller (DC-550), as follows:

− IC-600 computer - The primary component of the Automatic FlightControl System (AFCS). Controls the symbol generator, monitors,flight director and autopilot. Only the IC-600 #1 incorporates theautopilot functions.

− FLIGHT GUIDANCE CONTROLLER (GC-550) - Consists of a panelthat allows control of both Flight Director systems and autopilotfunctions. The GC-550 provides means for engaging the autopilotand the yaw damper, selecting the flight director modes and the flightdirector coupling. The Flight Guidance Controller also provides themeans for the remote selection of course, heading, vertical speedtarget, indicated airspeed target, Mach targets and preselectedaltitude.

− PITCH AND TURN CONTROLLER (PC-400) - Consists of a panelwith a Turn Control Knob and a Pitch Control Wheel. These controlsallow the pilot to manually maneuver the airplane with the autopilotengaged.

− DISPLAY CONTROLLER PANEL (DC-550) - The DC is used toselect various features on the PFD. These include HorizontalSituation Indicator (HSI) formats, navigation sources, weather displayand bearing pointer selection.

The Automatic Flight Control System interfaces with the followingsystems:

− INERTIAL REFERENCE SYSTEM (IRS): provides pitch, roll andacceleration information to the autopilot system via IC-600-1.

− RADIO MANAGEMENT SYSTEM: provides navigation data to theIC-600, including short range navigation data, VOR bearings, ILSapproach data, marker beacon tone detection and transmission,DME features and ADF.

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− AIR DATA COMPUTERS (ADCs): supply pressure altitude,barometrically corrected altitude, true airspeed, calibrated airspeed,vertical speed, Mach number, static air temperature and total airtemperature to both IC-600.

− RADIO ALTIMETER SYSTEM: provides radio altitude, low altitudeawareness and decision height information on the PFD.

− STALL PROTECTION SYSTEM: provides sensitive, visual and auralindications of an impending stall. If a stall condition is near to occur,the system actuates the stick shaker, disengages the autopilot and, ifnecessary, actuates the stick pusher.

− ENHANCED GROUND PROXIMITY WARNING SYSTEM(EGPWS/GPWS): receives, from IC-600-1, the glideslope deviation,localizer deviation, selected decision height, selected course, packeddiscrete and selected terrain range.

− ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS): presentinformation to the flight crew. Consists of two Primary Flight Displays(PFD), two multi function displays (MFD) and one EICAS display.

− HORIZONTAL STABILIZER CONTROL UNIT (HSCU): provides, toboth IC-600 #1 and #2, the horizontal stabilizer position. It alsoreceives, from IC-600, the autopilot command, when the autopilot isengaged, and the amount of trim demanded.

− AURAL WARNING UNIT (AWU): receives signal from the autopilot,generates the appropriate messages and tones and send the audiosignal to the Audio Digital System, which routes the messages to thespeakers.

− FLAP ELECTRONIC CONTROL UNIT (FECU): moves the inboardand outboard flap panels and sends flap position signal to theautopilot system.

− FLIGHT MANAGEMENT SYSTEM (FMS): provides high accuracy inlong range lateral navigation.


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FLIGHT GUIDANCE SYSTEM

The Flight Guidance System may perform three separate functions:the Flight Director, Autopilot and Autopilot Monitoring.

FLIGHT DIRECTOR

The Flight Director function provides pitch and roll attitude commandsbased on data from a variety of sensors, including attitude, heading, airdata, radio altimeter, navigation and pilot inputs. These attitudecommands are sent to the PFD for pilot display, to the autopilot forautomatic airplane control and to the autopilot monitors.

AUTOPILOT

The autopilot provides yaw stabilization and follows pitch and rollattitude commands from the flight director.

The autopilot/yaw damper monitors continuously check autopilotfunctions and operation. In case of failure, they are capable ofdisengaging the autopilot and yaw damper, independent of theautopilot processor hardware.

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FLIGHT DIRECTOR MODES

Flight Director mode selection is accomplished through the FlightGuidance Controller. Each mode selector button is illuminated for thearmed and captured mode. Also, each active mode is annunciated onthe PFD display and this annunciation makes the distinction betweenarmed and captured modes. The various modes may be divided intotwo categories: Lateral and Vertical modes.

LATERAL MODES

Lateral modes are those modes related to heading or roll control. Theynormally provide commands based on navigation sources.

HEADING HOLD MODE

Heading Hold mode is the default Flight Director mode when no otherlateral mode is selected. The Heading Hold mode provides rollcommands to maintain the heading at the moment of modeengagement. Once this mode is selected, the heading reference isestablished one second after the system detects a bank angle of lessthan 6º. A bank angle command of zero degrees is used (wings level)until the heading reference is established.

The ROL green label is displayed on the PFD to indicate the mode isengaged. Only the pilot’s side primary heading is used by this mode. Ifthis data is invalid, the Wings Level submode is used.

The Heading Hold mode is divided into Roll Hold submode, Turn Knobsubmode and Wings Level submode.

ROLL HOLD SUBMODE

The Roll Hold submode is entered from Heading Hold mode, with theautopilot engaged, by using the Touch Control Steering Button (TCS)to manually fly the airplane to a bank angle greater than 6°. Thesystem maintains the bank angle at the time the TCS button isreleased. Roll Hold submode may be canceled by either manuallyflying the airplane to less than 6° with the TCS button, by moving theTurn Control Knob out of detent or by selecting another lateral mode.This mode is annunciated on the PFD by the ROL green label.


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TURN KNOB SUBMODE

The Turn Knob submode allows the pilot to generate a roll attitudecommand manually with the Turn Control Knob. Moving the TurnControl Knob out of detent, with the autopilot engaged, cancels allother lateral modes including Heading Hold mode in both FlightDirectors.

When the Turn Control Knob is out of detent, the autopilot will maintaina roll attitude proportional to the displacement of the knob. Theautopilot will revert back to Heading Hold mode when the turn knob isplaced in the detent position. Turn Knob submode is annunciated onthe PFD by the ROL green label when out of detent and the autopilot isengaged. When the autopilot is disengaged and the Turn Control Knobis out of detent, the TKNB label is displayed in the PFD and theautopilot engagement is inhibited.

WINGS LEVEL SUBMODE

The Wings Level submode provides a roll command of 0º. This modeis active in the Go Around mode, Windshear mode or if the primaryheading data is invalid. Therefore, this mode is available even if eitherattitude source is invalid. This mode is annunciated on the PFD by theROL green label.

HEADING SELECT MODE (HDG)

The HDG mode is selected by pressing the HDG button on the flightguidance controller or by arming LOC, VOR, VAPP, or BC. This modeallows the Flight Director to track the EHSI heading bug, as set by theheading select knob. The Heading Select mode is annunciated on thePFD by the green HDG label.

The mode will be inhibited by the following conditions:

− Turn Control Knob out of detent with autopilot engaged.− Displayed heading invalid.

The mode will be canceled if any of the following conditions occur:

− Pressing the HDG button.− Changing the displayed heading source on the PFD.− LOC & BC mode capture.− VOR & VAPP capture.− Pressing the Go Around button.

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LOW BANK MODE

The Low Bank mode allows the pilot to select reduced bank angle forthe HDG mode. Bank angle limit will be reduced from 27° to 14°whenever this mode is active. The mode is selected by pressing theBNK button in the Flight Guidance Controller. This mode isannunciated only while the Heading Select mode is active, but remainsselected if Heading Select mode is deactivated, being reactivated andannunciated if Heading mode is selected again. The Low Bank mode isautomatically selected when climbing above 25000 ft and automaticallycanceled when descending below 24750 ft.

VOR NAV MODE (VOR)

The VOR NAV mode allows automatic capture and tracking of bothinbound and outbound VOR radials. The VOR mode is selected bypressing the NAV button in the Flight Guidance Controller, with VORselected on the PFD. Upon selection of VOR NAV mode, the HDGselect mode will automatically be engaged. This triggers the greenHDG annunciation on the PFD in conjunction with an armed white VORNAV annunciation, also on the PFD.

At the proper time, based on course error and beam deviation, thecapture of VOR mode will cancel the HDG selected mode.

The mode will be canceled or inhibited if any of the following conditionsoccur:− Pressing the NAV button.− Selecting VAPP or HDG modes.− Changing the displayed NAV source on the PFD.− Changing the displayed heading source on the PFD.− When the displayed heading is invalid.− When the displayed NAV source is invalid for more than 5 seconds.− Pressing the Go Around Button.− Turn Control Knob out of detent with autopilot engaged.


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VOR APPROACH MODE (VAPP)

The VOR Approach mode provides the same capabilities as the VORNAV mode, with higher gain for operation close to the station.

It is recommended to select VAPP mode only on the final approachsegment. Therefore, the outbound segment should be flown usingsome other mode.

This mode is selected by pressing the APR button on the FlightGuidance Controller, with VOR displayed on the PFD. This mode iscanceled or inhibited by the same conditions as the VOR NAV mode.

Selecting VOR Approach mode, the HDG select mode willautomatically be engaged providing the green HDG annunciation onthe PFD in conjunction with the armed VOR approach and white NAVannunciation, also on the PFD.

LOCALIZER MODES (LOC/BC)

The Localizer Modes allow automatic capture and tracking of localizertransmitters. Both front course (LOC) and back course (BC)approaches are supported.

The back course approach operates similar to the front courseapproach, except that the beam deviation is inverted, allowing thesystem to approach the runway 180° from the front-course.

Select the Localizer mode by pressing the NAV or APR buttons on theflight guidance controller with ILS as the selected navigation source. Inthis case, the HDG select mode is automatically selected and thelocalizer is armed. On an ILS approach, when the localizer is armedand the APR button is pressed, the Glide Slope is also armed.

The localizer mode captures are based on course error and beamdeviation. At the point of capture, the current armed mode (LOC or BC)is selected and locked, while HDG select mode is canceled. The LOCmode capture or BC mode capture is annunciated on the PFD by agreen LOC or green BC label, respectively.

After captured, the mode will be canceled or inhibited if any of thefollowing conditions occur:

− Pressing the NAV or APR buttons.− Selecting HDG mode.− Changing the displayed NAV source on the PFD.− Changing the displayed heading source on the PFD.− When the displayed heading is invalid.

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− When the displayed NAV source is invalid for more than 5 seconds.− When the displayed Glide Slope deviation is invalid for more than 5

seconds, with GS mode captured.− When the on-side attitude is invalid.− When the selected air data source is invalid.− Pressing Go Around button.− Turn Control Knob out of detent with autopilot engaged.

LNAV MODE

The LNAV mode allows the Flight Director to capture and track the rollsteering signal from the long range navigation system (FMS/GPS).

With FMS selected on the PFD, select LNAV mode by pressing theNAV button on the Flight Guidance Controller. This mode willautomatically engage HDG select mode, triggering a green HDGannunciation on the PFD in conjunction with a white LNAVannunciation, also on the PFD.

The mode will be canceled or inhibited if any of the following conditionsoccur:

− Pressing the NAV button.− Selecting HDG mode.− Changing the displayed NAV source on the PFD.− Changing the displayed heading source on the PFD.− When the displayed heading is invalid.− When the lateral steering command is invalid.− Pressing the Go Around button.− Turn Control Knob out of detent with autopilot engaged.


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VERTICAL MODES

Vertical modes are those modes related to pitch control. Due to thenecessity of maintaining the wings leveled during Go Around, thisvertical maneuver may also be considered as a lateral mode.

PITCH HOLD MODE

The Pitch Hold mode is the default mode that controls the airplanewhen no other Flight Director mode is selected.

The Pitch Hold mode is synchronized to the existing pitch attitude andprovides an error signal to the command bars and autopilot function.

By pressing the Touch Control Steering Button (TCS), the pilot maymanually change the pitch attitude and then allow the system toresynchronize to the new attitude when the button is released.

Should the autopilot be engaged and the Flight Director is in the pitchhold mode, pitch attitude reference can be changed by rotating thepitch control wheel on the pitch and turn controller.

The pitch control wheel allows continuous variable rates andamplitudes of the pitch reference. A PIT label is displayed on the PFDto indicate mode engaged.

ALTITUDE HOLD MODE (ALT)

The Altitude Hold mode generates an altitude error signal from areference altitude and provides a pitch command, which allows theautopilot to maintain altitude.

The Altitude Hold mode is selected by pressing the ALT button on theFlight Guidance Controller or can also be activated automatically by thealtitude preselect mode. This mode is annunciated on the PFD by theALT label.


− Pressing the ALT button.− Selecting VS, FLC, or SPD modes.− Glide slope capture.− When the air data is invalid.− Pressing the Go Around Button.− Pitch control wheel moved with autopilot engaged.

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ALTITUDE PRESELECT MODE (ASEL)

The Altitude Preselect mode provides means for the system to climb ordescend to a predetermined altitude and then level off and maintainthe preselected altitude.

Preselected altitude is set through the ASEL knob on the FlightGuidance Controller and is displayed on the top right corner of thePFD. This mode is annunciated by the white ASEL label on the PFD.

Pitch Hold, Speed Hold or Vertical Speed Hold must be used to climbor descend towards the preselected altitude or Flight Level Change(FLC).

The ASEL mode will arm automatically if the airplane climbs ordescends towards a preselected altitude. The ASEL mode willautomatically capture and cancel any existing mode at the appropriatepoint based on preselected altitude error and vertical speed. Thesystem will automatically switch to altitude hold mode after the airplanehas leveled off at the selected altitude.

The mode will be canceled and/or inhibited if any of the followingconditions occur:

− Changing the preselected altitude.− Selecting ALT, VS, FLC, or SPD modes.− Glide slope capture.− When the air data is invalid.− Pressing the Go Around Button.

FLIGHT LEVEL CHANGE MODE (FLC)

The Flight Level Change mode (FLC) provides means of climbing ordescending to a preselected altitude at a pre-programmed schedule.

When the preselected altitude is above the current altitude and theflight level change mode is selected, the Flight Director provides aspeed command at the predetermined climb speed schedule. Whenthe preselected altitude is below the current altitude and FLC isselected, the FD provides a command to descend at a determined rateof descent. The PFD will display the current IAS, Mach or verticalspeed bug as appropriate and the target speed can be adjusted onlyby deselecting the flight level change mode.

As the airplane approaches the preselected altitude, the Flight Directorwill cycle among ASEL ARM, ASEL CAP, and ALT HOLD to capturethe preselected altitude.


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The following protections are provided with this mode:

− Maximum normal and longitudinal acceleration: 0.1 G.− Maximum airspeed: VMO or MMO.

− System will maintain the preselected altitude.

The Flight Level Change mode may be activated by selecting analtitude and pressing the FLC button in the Flight Guidance Controller.This mode is annunciated on the PFD by the CLB label, when followingthe IAS/MACH climb profile, or by the DES label when following avertical descent profile of - 2000 ft/min.


− Pressing the FLC button.− Changing the preselected altitude.− Selecting ALT, VS, FLC, or SPD modes.− Glide slope capture.− When the air data is invalid.− Pressing the Go Around Button.

DESCENT RATE SCHEDULE:

For all EICAS versions:

From 41000 ft to 12000 ft, the descent rate schedule is −2000 ft/min.

From 12000 ft to 10000 ft the descent rate schedule is −2000 ft/min to−1000 ft/min.

From 10000 ft and below the descent rate schedule is −1000 ft/min.

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CLIMB RATE SCHEDULE:

For airplanes equipped with EICAS versions up to 20.5, the climb rateschedule is presented in the chart below:

150

170

190

210

230

250

270

290

310

0 3000 6000 9000 12000 15000 18000 21000 24000 27000 30000 33000 36000 39000

ALTITUDE - ft

IND

ICA

TE

D A

IRS

PE

ED

- k

t

M = 0.65

IAS = 290 kt

IAS = 240 kt

135B

JAO

M03

A -

12N

OV

2004

10000 12000 21600 39000

240

290

197

For airplanes equipped with EICAS versions 20.6 and on, the climbrate schedule is presented in the chart below:

150

160

170

180

190

200

210

220

230

240

250

260

270

280

290

300

0 3000 6000 9000 12000 15000 18000 21000 24000 27000 30000 33000 36000 39000

ALTITUDE - ft

IND

ICA

TE

D A

IRS

PE

ED

- k

t

IAS = 270 kt

IAS = 240 kt

M = 0.65

8000 14000 25051 39000 41000

240

270

197

188

13

5B

JAO

M0

3B

- 1

2F

EV

20

05


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SPEED HOLD MODE (SPD)

The Speed Hold mode is used to maintain airspeed or Mach numberwhile flying to a new altitude. Indicated airspeed should be used below25000 ft and Mach number above 25100 ft.

The Speed Hold mode is also designed to provide overspeed andunderspeed protections.

Speed hold mode is selected by pressing the SPD button on the FlightGuidance Controller. This mode is annunciated on the PFD by the SPDlabel, when maintaining IAS, or by the MACH label when maintainingMach number. Selection of Speed Hold mode cancels other verticalmodes, except the altitude preselect arm mode and Glide Slope armmode.

Speed Hold mode is automatically selected when the FLC button ispressed and the preselected altitude is above the current altitude.

Different Speed Target can be selected by using the Speed Set knobin the Flight Guidance Controller. Pressing the SPD knob allows thepilot to toggle between IAS target and MACH target to set airspeed.


− Maximum normal acceleration: 0.1 G.− Maximum normal acceleration on entering overspeed: 0.3 G.− Maximum airspeed: VMO or MMO.− Minimum airspeed: Shaker actuation speed.− System will maintain the preselected altitude and airspeed.


− Pressing the SPD button.− Selecting ALT, VS, or FLC modes.− Altitude preselect capture.− Glide slope capture.− When air data is invalid.− Pressing the Go Around Button.

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VERTICAL SPEED HOLD MODE (VS)

The Vertical Speed hold mode is used to maintain or to make changesto the vertical speed. The Vertical Speed hold mode ranges from- 6000 to + 6000 ft/min, with a resolution of 100 ft/min.

The Vertical Speed Hold mode is selected by pressing the VS buttonon the Flight Guidance Controller or automatically, when FLC button ispressed and the preselected altitude is below the current altitude. Thismode is annunciated on the PFD by the VS label.

Selection of this mode cancels other vertical modes, except the altitudepreselect arm and Glide Slope arm.

Vertical speed may be changed by using the Speed Set knob, on theflight guidance controller.


− Maximum airspeed: VMO.− Minimum airspeed: Shaker actuation speed.


− Pressing the VS button.− Selecting ALT, SPD, or FLC modes.− Altitude preselect capture.− Glide slope capture.− When air data is invalid.− Pressing the Go Around Button.


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GLIDE SLOPE MODE (GS)

The Glide Slope mode allows automatic capture and tracking to GlideSlope transmitters. Select Glide Slope mode by pressing the APRbutton with ILS as a navigation source.

Selecting Glide Slope mode automatically arms GS (in conjuction withLOC). The PFD will display a white localizer LOC and a white GlideSlope GS annunciation. The localizer mode capture will occur with agreen LOC annunciation on the PFD. The Glide Slope mode capture,with a green GS annunciation on the PFD, will occur only afterLocalizer mode has been captured.

After captured, the GS mode will be canceled or inhibited if any of thefollowing conditions occur:

− Pressing the APR or NAV buttons.− Lost Localizer mode.− Selecting ALT, SPD, VS, or FLC modes.− Glide slope deviation invalid for a period greater than 5 seconds.− Pressing the Go Around Button.

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GO AROUND MODE

TAKEOFF SUBMODE

The Takeoff submode provides a wings level command and a fixedpitch up attitude command of 14° (for flaps at 9°), which is indicated bythe Flight Director command bars on the EADI.

This mode is selected by pressing any of the Go Around buttons on thethrust levers and annunciated by the ROL label and TO label, both onthe PFD.

The Takeoff submode will be canceled if any of the following conditionsoccur:

− Pushing the TCS button.− Selecting ALT, SPD, VS, or FLC mode.− Transition to capture Altitude Preselect mode.− Air data computer source selection is changed.

The Takeoff submode is available on the ground with airspeed below60 KIAS or in flight within 400 ft above the runway.

The Go Around mode, as well as the Vertical Speed Control knob, willbe inhibited while Takeoff submode is engaged.

After reaching the 400 ft delta, pressing the Go Around button willengage the Go Around mode. Once the 400 ft boundary is crossed, the400 ft delta requirement will be ignored, to avoid restricting any GAmaneuvers later in the flight.

If the autopilot is selected with the Takeoff submode engaged, thissubmode will drop into Pitch Hold mode and synchronize to the currentattitude. The Takeoff submode will not be coupled to the autopilot,which may be used after climbing above the airplane MinimumEngagement Height (MEH).

When the autopilot is not engaged, wings level will be the active lateralmode and the ROL label will be displayed on the PFD.

A Pitch Limit Indicator (PLI) is displayed on the EADI sphere when themargin prior to the stick shaker set point is below or equal to 10°. In thecase of an invalid Stall Protection Computer signal, the PLI will bebiased out of view and an amber AOA annunciation will be displayedon the PFD.


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01 REVISION 7

GO AROUND SUBMODE

The Go Around Submode should be selected once the decision todiscontinue the approach has been taken. Although commanding anose up attitude, the need to maintain wings leveled causes this modeto incorporate both lateral and vertical modes features.

- Speed Target Submode:

The Speed Target submode will command airplane pitch in orderto allow a climbing turn at an airspeed of around 1.23 VS. Once apositive rate of climb has been achieved, the Speed Targetsubmode will limit the pitch angle at 10° nose up. The systemmanages airspeed, altitude and comfort. Therefore, accelerationsare limited to avoid passenger discomfort, while maintaining targetairspeed. If the airspeed can not be maintained, altitude will beheld.

The Speed Target mode will initially command the Flight DirectorCommand Bar and the autopilot pitch up attitude to 10° nose upfor at least 20 seconds. After this, the Flight Director provides apitch up command based on the IAS Speed Hold mode followingthe go-around speed preselected on the airspeed bug and limitedwithin 1.23 VS and 170 KIAS.

NOTE: The Flight Director will revert automatically to IAS speedhold, without waiting 20 seconds if at the time the GoAround button is pressed or during the time the GoAround mode is engaged, the airplane is below 1.23 VS.

The airspeed bug is displayed on the airspeed tape on the PFDand a pitch limit indicator is displayed on the EADI. If the StallProtection Computer signal becomes invalid, the PLI is removed.

The mode may be engaged by pressing any of the Go Aroundbuttons on the thrust levers. The submode may be engaged onlyat radio altitudes below 2500 ft, or below 15000 ft pressurealtitude for an invalid Radio Altimeter signal. This feature isprovided to protect against inadvertent Go Around selectionsduring cruise.

The autopilot may be coupled to the Speed Target submodeabove the airplane’s Minimum Use Height (MUH), but will not beinhibited below the MUH.

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AUTOPILOT


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The GA label is annunciated on the PFD during the first 20seconds, when the 10° pitch up command exists. When the IASpreselected speed bug is used on the go-around, the GA labelswitches to the IAS label and the system provides the pitchcommand based on the IAS Hold mode.

The Speed Target submode will disengage on selection of a newvertical mode. The submode will ignore a preselected altitudebelow the airplane and will not fly away from a preselected altitudeabove the airplane. Altitude Preselect mode will be inhibited if thepreselect altitude is less than Speed Target submodeengagement altitude plus 400 ft (pressure altitude). This feature isprovided to avoid the airplane leveling off if the pilot has notreadjusted the preselected altitude to the new missed approachaltitude.

The Speed knob will be inhibited while GA mode is engaged.

When the autopilot is not engaged, wings level will be the activelateral mode and the ROL label will be displayed on the PFD. Ifthe autopilot is engaged, the lateral mode will remain wings leveland will also be displayed as ROL on the PFD.

WINDSHEAR ESCAPE GUIDANCE MODE

The Windshear Escape Guidance mode is provided in order to recoverfrom a windshear situation.

For further information on windshear detection and escape guidancesystem, refer to Section 2-4 – Crew Awareness.


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01 REVISION 7

AUTOPILOT DISENGAGEMENTThe autopilot is normally disengaged through the AutopilotEngage/Disengage button or through the quick disconnect button onthe control wheel.

A voice message AUTOPILOT is generated when the autopilot isdisengaged.

This message is presented at any altitude in case of intentionaldisengagement or due to an autopilot failure and may be canceledaccording to the following associated conditions:

Associated Conditions Cancellation

Above 2500 ft radio altitude witha valid Radio Altimeter signal.

Self canceled.

Below 2500 ft radio altitude witha valid Radio Altimeter signal.

Pressing the Autopilot QuickDisconnect Button twice.

Invalid Radio Altimeter signal. Pressing the Autopilot QuickDisconnect Button twice.

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AUTOPILOT


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WARNING AUTOPILOT FAIL Autopilot has failed and hasbeen automatically disengaged.

AUTO TRIM FAIL Automatic pitch trim has failed.

AP ELEV MISTRIM A pitch mistrim condition exists.

CAUTION AP AIL MISTRIM A roll mistrim condition exists.

LATERAL MODE OFF Inadvertent loss of the LateralFlight Director mode.

VERTICAL MODE OFF Inadvertent loss of the VerticalFlight Director mode.

YAW DAMPER FAIL Yaw Damper has failed and hasbeen automatically disengaged.


FLIGHT GUIDANCE CONTROLLER

NOTE: All the mode selector buttons described below are illuminatedto indicate whether the associated mode is armed or captured.

1 - FLIGHT DIRECTOR BUTTON− Allows the Flight Director bars to be displayed on the associated

PFD.

2 - LATERAL MODE SELECTOR BUTTONS− Select lateral operating modes of the autoflight system, as

follows:− HDG: selects heading hold and heading select modes.− NAV: selects VOR NAV mode and allows selection of LOC/BC and LNAV modes.− APR: selects VOR approach mode and allows selection of

LOC/BC and GS modes.− BNK: selects Low Bank submode.

3 - AUTOPILOT ENGAGE BUTTON− Pressed once engages the autopilot and the yaw damper.

Pressed again, disengages the autopilot only, keeping the yawdamper engaged.


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4 - VERTICAL MODE SELECTOR BUTTONS− Select vertical operating modes of the autoflight system, as

follows:− SPD: selects Speed Hold mode.− FLC: selects Flight Level Change mode.− VS: selects Vertical Speed hold mode.− ALT: selects Altitude Hold mode.

5 - ALTITUDE PRESELECT KNOB− Allows preselection of altitude in 100 ft increments.

6 - COURSE SELECTOR KNOB− Allows selection of course in 1° increments.− Pressing the knob synchronizes the selected course to the VOR

bearing.

7 - VERTICAL SPEED CONTROL KNOB AND IAS/M SELECTORBUTTON− Pressing the knob toggles between the speed modes MACH

and IAS.− When in SPD mode, rotation of this knob allows selection of

indicated airspeed in one-knot increments or Mach Number in0.01 increments.

− When in VS mode, rotation of this knob allows selection ofvertical speed in 100 ft/min increments.

8 - YAW DAMPER ENGAGE BUTTON− Pressed once, engages only the Yaw Damper. Pressed again

disengages the yaw damper and the autopilot, if it is engaged.

9 - AUTOPILOT COUPLE BUTTON− Allows the pilot’s or copilot’s Flight Director commands to control

the autopilot. The couple button can be pressed with theautopilot engaged or disengaged. However, if the Flight Directoris switched, the modes will drop out and the autopilot will remainengaged (if already engaged) and revert to basic autopilot mode(pitch and roll).

10 - HEADING SELECT KNOB− Allows selection of heading in 1° increments.− Pressing this knob synchronizes the heading selection to the

current displayed heading.

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FLIGHT GUIDANCE CONTROLLER


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PITCH AND TURN CONTROLLER

1 - PITCH CONTROL WHEEL− Manually controls the pitch when the autopilot is engaged and

the Pitch Hold mode is selected.− Pitch control wheel operation is inhibited if any vertical mode,

except the Pitch Hold mode, is selected in the Flight Director.

2 - TURN CONTROL KNOB− Manually controls the roll attitude when the autopilot is engaged.− The control has a center detent position at the wings leveled

position. The control remains at the current position whenreleased.

PITCH AND TURN CONTROLLER

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CONTROL WHEEL

1 - TOUCH CONTROL STEERING BUTTON (TCS)

− Allows manual maneuvering of the airplane without disengagingthe autopilot. The airplane may be maneuvered to any desiredpitch attitude while the TCS button is pressed. When the buttonis released, the following occurs:− Primary servos reengage.− The computer synchronizes itself to the new pitch attitude

and vertical mode and maintain it.− Lateral control is returned to the previously selected lateral

mode (return to the lateral mode is filtered to prevent rapidmaneuvers).

− After glide slope capture in APR mode with the autopilotengaged, if the TCS button is pressed and released, theautopilot will resume the controls and turn the airplane to the ILScenter beam.

2 - QUICK DISCONNECT BUTTON

− Provides the means to disengage autopilot and yaw damper.− The pilot’s and copilot’s buttons are interconnected to allow

autopilot cancellation from either seat.− In case of the autopilot is disengaged and the button is pressed,

the voice message AUTOPILOT will be canceled in 2 seconds.


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CONTROL WHEEL

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THRUST LEVERS

1 - GO AROUND BUTTON

− Selects the Go Around mode (Takeoff submode, Go AroundSpeed Target submode and Windshear mode).

− The button also forces the Flight Director into either the GoAround mode or the Windshear mode, depending on thewindshear signal.

THRUST LEVERS


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DISPLAY CONTROLLER (DC-550)

1 - NAVIGATION SOURCES SELECTOR BUTTON

− Provides the selection of the VHF NAV (VOR, ILS and MLS) asnavigation source for the EHSI. If the VHF NAV is alreadyselected, pressing the NAV Button selects the opposite VHFNAV as navigation source for the on-side EHSI. Pressing theNAV Button once again will restore the normal operation: VHFNAV 1 information presented on the PFD 1 and VHF NAV 2information presented on the PFD 2.

2 - FMS SOURCE SELECTOR BUTTON (optional)

− Provides the selection of the FMS as navigation source for theEHSI.

− On airplanes equipped with dual FMS, pressing the FMS Buttonfor the second time selects the opposite FMS as navigationsource for the on-side EHSI (and for the on-side MFD MAP).Pressing the FMS Button once again will restore the normaloperation: FMS 1 information presented on the PFD 1 (and MFD1) and FMS 2 information presented on the PFD 2 (and MFD 2).

3 - BEARING SELECTOR KNOB

OFF: The associated PFD bearing pointers are disabled.NAV 1 (2): Selects the respective VHF NAV as source for the

associated bearing pointer.ADF: Selects the respective ADF as source for the

associated bearing pointer.FMS: Selects the FMS as source for the associated bearing

pointer.

4 - DECISION HEIGHT SETTING AND IC-600 TEST KNOB

− Provides the Radio Altimeter (RA) decision height setting.− When pressed on ground provides the IC-600 and RA test

activation. Refer to Section 2-4 – Crew Awareness for furtherinformation on test function and Section 2-17 – FlightInstruments for further information on decision height settingand RA test in flight.

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DISPLAY CONTROLLER PANEL (DC-550)


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PFD INDICATORS

1 - ARMED LATERAL MODE (white)− − Indicates which lateral mode is armed.− − The mode annunciation is removed if the Flight Director fails.

2 - CAPTURED LATERAL MODE (green)− − Indicates which lateral mode is captured.− An amber FD FAIL is displayed in this field to indicate Flight

Director failed.− − The mode annunciation is removed if the Flight Director fails.

3 - AUTOPILOT MESSAGE FIELD− Indicates autopilot status.− Messages are mutually exclusive and therefore only one

message can be displayed at a time.− The following messages may be displayed:

MESSAGE COLOR MEANINGAP Autopilot engaged.

AP TEST Green Autopilot test mode is active immediatelyafter power up.

TCS TCS submode is engaged (autopilot isengaged).

TKNB AmberTurn control knob is out of detentposition (autopilot is disengaged).

AP

When the autopilot is normallydisengaged, the green AP annunciationturns amber and flashes for 5 seconds,then becomes steady.

AP Red

If the autopilot is engaged and a failureoccurs, the green AP annunciation turnsred and flashes for 5 seconds, thenbecomes steady. The AP annunciationappears in conjunction with theAUTOPILOT FAIL message on theEICAS and is removed when theautopilot is disengaged through theQuick Disconnect Button.

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4 - FLIGHT DIRECTOR COUPLE ARROW− Indicates which Flight Director the autopilot is coupled to.− The mode annunciation is removed if the Flight Director fails.

5 - YAW DAMPER ENGAGED ANNUNCIATION− Color:

− Green: indicates the yaw damper is engaged.− Amber: when the yaw damper is normally disengaged the

annunciation flashes for 5 seconds and then extinguishesitself.If the yaw damper is engaged and a failure occurs, theannunciation flashes for 5 seconds then becomes steadyuntil it is disengaged through the Quick Disconnect Button.

6 - CAPTURED VERTICAL MODE (green)− Indicates which vertical mode is captured.− The mode annunciation is removed if the Flight Director fails.

7 - MODE TRANSITION ANNUNCIATOR− Each transition is annunciated by a box around the mode that is

being transitioned. The box will highlight the new mode for5 seconds and then disappear.

8 - ARMED VERTICAL MODE (white)− Indicates which vertical mode is armed.− The mode annunciation is removed if Flight Director fails.

9 - ALTITUDE PRESELECT DISPLAY

− Ranges from – 900 to 45000 ft with a resolution of 100 ft.− The digits and bug are cyan and the box is white. They become

amber 1000 ft prior to reaching the preselected altitude. Oncethe airplane is within 250 ft of the preselected altitude, the boxreturned to white. If the airplane exceeds the preselectedaltitude by more than 250 ft, the box turns amber.

− Large digits display hundreds, thousands and tens of thousands.Smaller digits, which are always zeros, display tens and ones.

− The bug moves according to the digital altitude preselect value.− If the preselected altitude value is not within the displayed range

of the altitude scale, the bug will stay at the respective end ofscale, half-visible and unfilled.


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10 - COMMAND BAR AND AIRPLANE SYMBOL

− Color: magenta.− Indicates pitch and roll Flight Director commands.− Command bar is removed if the Flight Director fails or if the

opposite side Flight Director selected source or tuned frequencyis different.NOTE: The command bar and airplane symbol may be

presented in either V-bar or cross-bar formats,depending on operator selection.

11 - SELECTED HEADING BUG

− Color: magenta.− Displayed full time on the PFD, unless when the PFD is in arc

format.− When setting the selected heading value, the bug will move

around the heading scale.

12 - VERTICAL SPEED TARGET DISPLAY

− Color: cyan.− Ranges from 0 to 9900 ft/min with a resolution of 100 ft/min.− Displayed only when Vertical Speed Hold mode is selected in

either Flight Director.

13 - SELECTED HEADING DIGITAL READOUT− Color:

− Digits: cyan.− Label: white.

− Indicates the heading selected through the Flight GuidanceController panel.

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14 - GS/LOC/ILS COMPARISON MONITOR DISPLAYS

− Label: GS, LOC or ILS.− Color: amber.− Glide Slope comparison monitor (GS label) is displayed while in

GS CAP and below 2500 ft if there is a difference of 0.7 dotdeviation between the PFDs indication. If the radio altitudeoutput is invalid, the monitor will then be activated by GS CAPonly.

− Localizer comparison monitor (LOC label) is displayed while inapproach mode, below 2500 ft if there is a difference of 0.5 dotdeviation between the PFDs indication. If the radio altitudeoutput is invalid, the monitor will then be activated by GS CAPonly.

− ILS comparison monitor display is annunciated when both GSand LOC comparison monitors are displayed simultaneously.

15 - AOA INDICATION

− Color: amber.− Indicates loss of PLI indication due to an invalid Stall Protection

Computer signal.

16 - OVERSPEED/UNDERSPEED WARNING DISPLAY

− Color: amber.− Label: MAX SPD for overspeed condition.

MIN SPD for underspeed condition.− Activated by the Flight Director.− Remains displayed as long as the condition exists.

17 - INDICATED AIRSPEED/MACH TARGET DISPLAY

− Color: digits are cyan and box is white.− Ranges from 80 KIAS to VMO with a resolution of 1 KIAS or from

0.2 Mach to MMO with a resolution of 0.01 Mach.− Displayed full time.− Bug moves according to the indicated airspeed/Mach target

value set.− If the indicated airspeed/Mach value is not within the displayed

range of the airspeed scale, the bug will stay at the respectiveend of the scale, half-visible and unfilled.


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PFD INDICATORS(CROSS-BAR FORMAT)

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PFD INDICATORS(V-BAR FORMAT)


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EICAS INDICATORS

1 - ROLL MISTRIM ANNUNCIATION

− Color: amber.− − Indicates that a roll mistrim exists, which may cause an abrupt

roll command at the time the autopilot is disengaged.− − Direction of arrow indicates the side the roll trim must be

commanded to eliminate the condition.− − It is displayed in conjunction with the AP AIL MISTRIM message

on the EICAS.

ROLL MISTRIM ANNUNCIATION

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CATEGORY II APPROACH (OPTIONAL)The IC-600 may be optionally equipped with a Category II checklistlogic warning which is automatically activated whenever the DecisionHeight is selected between 80 and 200 ft through the RA knob on theDisplay Control Panel.

CATEGORY II CONDITIONS

The required conditions to obtain a Cat II valid conditions are:

− Radio altitude between 2500 and 80 ft.− Flaps 22°.− NAV 1 on pilot’s side and NAV 2 on copilot’s side, both tuned to

the same frequency.− An active approach mode selected.− Both Flight Directors operational (command bars visible).− Attitude and heading valid on both PFDs.− Glide slope and localizer deviation valid on both PFDs.− No reversions (SG, IRS or ADC) modes selected on both PFDs.− Valid airspeed and barometric altitude on both PFDs.− No comparison monitors are tripped (attitude, heading,

airspeed, barometric altitude, localizer, glide slope and radioaltitude) on both PFDs.

− No back course selected.− Autopilot engaged.− Cat II Decision Height setting on both Display Control Panels

(greater than 80 ft and less than 200 ft).

NOTE: CAT II approaches are allowed using either the Autopilot orFlight Director for guidance.

If all conditions are met, a green CAT 2 annunciation is displayed onthe PFDs. If any of the required conditions for establishing CAT 2 goesinvalid, the green CAT 2 will be replaced by flashing amber CAT 2annunciation. It will flash for ten seconds and then go steady.


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LOCALIZER EXCESSIVE DEVIATION WARNING

The Localizer Excessive Deviation Warnings are active when CAT IIis valid. It is triggered between 500 ft and 80 ft of Radio Altimeterheight, when the ILS lateral deviation is greater than 1/3 dot. Thepilots will be alerted by:

− The lateral deviation bar on the EHSI are changed from green toamber.

− The lateral deviation scale changing from white to amber andflashing with a duty cycle of 0.5 second on followed by a 0.5second off.

NOTE: The on-side excessive deviation warning is also displayedwhen the cross-side system has detected an excessivedeviation.

GLIDESLOPE EXCESSIVE DEVIATION WARNING

The Glideslope Excessive Deviation Warnings are active whenCAT II is valid. It is triggered between 500 ft and 80 ft of RadioAltimeter height, when the ILS vertical deviation is greater than onedot or when the glideslope deviation warning is invalid for more thanfive seconds. The pilots will be alerted by:

− The GS pointer on the EADI changing from green to amber.− The GS scale on the EADI changing from white to amber and

flashing with a duty cycle of 0.5 second on followed by 0.5second off.

NOTE: The on-side excessive deviation warning is also displayedwhen the cross-side system has detected an excessivedeviation.


PFD INDICATORS

1 - CAT 2 ANNUNCIATION

− Indicates the Cat II condition.− Label: CAT 2.− Color:

− Normal condition: green.− Abnormal condition: amber.

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PFD INDICATORS


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SECTION 2-01

AIRPLANE DESCRIPTION

TABLE OF CONTENTSBlock Page

Introduction ........................................................................ 2-01-00 ..02

Airplane Description ........................................................... 2-01-00 ..03

Cockpit Arrangement ......................................................... 2-01-00 ..06

Interior Arrangement .......................................................... 2-01-00 ..07

Main/Glareshield Panels .................................................... 2-01-05 ..01

Control Pedestal................................................................. 2-01-05 ..02

Overhead Panel ................................................................. 2-01-10 ..01

Cockpit Partition ................................................................. 2-01-15 ..01

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INTRODUCTION

This Section is intended to present a general overview of the airplane,thus initiating the reader to the EMB-135BJ, which may, then, gothrough the Sections searching more detailed information on eachsystem.

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AIRPLANE DESCRIPTIONThe EMB-135BJ model is a low wing, T-tail pressurized airplane,powered by two high by-pass ratio rear mounted turbofan engines. Thetricycle landing gear is all retractable, with twin tires in each leg.

A glass cockpit panel has been developed with highly integrated on-board avionics, thus allowing pilots to better monitor airplane generaloperation.

There are three passenger cabin layout options, with front galley andrear toilet, permitting to carry up to 15 passengers. Convenientaccommodation is provided for the flight crew.

For detailed information on each system, refer to the appropriateSection in this manual.

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THREE VIEW DRAWING (EMB-135BJ)

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COCKPIT ARRANGEMENT

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INTERIOR ARRANGEMENTCROSS SECTION

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INTERIOR CONFIGURATION OPTION 1

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MAIN/GLARESHIELD PANELS (OPTION 1)

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CONTROL PEDESTAL (OPTION 1)

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MAIN/GLARESHIELD PANELS (OPTION 2)

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OVERHEAD PANEL

AIRPLANES UP TO S/N 854

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OVERHEAD PANEL

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COCKPIT PARTITION

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SECTION 2-09

AUXILIARY POWER UNIT

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-09-05 ..01Control System................................................................... 2-09-05 ..04APU Starting/Operation...................................................... 2-09-05 ..07EICAS Messages ............................................................... 2-09-05 ..08Controls and Indicators ...................................................... 2-09-05 ..09

APU Control Panel ......................................................... 2-09-05 ..09EICAS Indications........................................................... 2-09-05 ..10

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GENERALThe APU is a source of pneumatic and electrical power to be used eithersimultaneously with or independent of aircraft sources, while on theground or in flight. Basically, it is a constant-speed gas turbine engine,consisting of a single-stage centrifugal compressor, a reverse-flowannular combustion chamber, and a single-stage radial turbine.

The airplane is equipped with APU model T-62T-40C14, which iscontrolled by the Full Authority Digital Electronic Control (FADEC). Thecontrol system provides automatic, full-authority, fuel scheduling fromstart to full load operation, under all ambient conditions and operatingmodes. In addition, the FADEC automatically controls the APU to shutdown on the occurrence of certain failures or events during start oroperation.

An automatic APU shutdown may occur either on the ground or inflight, and takes place under the following conditions:

On the ground:

− fire− overtemperature− overspeed− underspeed− failure to start− failure to accelerate− failure to light− loss of speed data− external short− loss of FADEC signal− FADEC failure− bleed valve opening− low oil pressure− high oil temperature− oil pressure switch short− loss of EGT.

NOTE: In the event of fire, a 10 second delay is allowed before anautomatic APU shutdown is initiated.

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In flight:

− overspeed− underspeed− failure to start− failure to accelerate− failure to light− loss of speed data− external short− loss of FADEC signal− FADEC failure.

The APU compartment is located in the airplane’s tailcone, isolated bya titanium firewall. On the left side of the APU compartment, aninspection door allows access and inspection of the APU’scomponents.

The APU starter-generator shaft drives an air-cooling fan. Air is drawnthrough a NACA air inlet located on the left side of the tailcone. APUdraining is ducted to the airplane skin on the right side of the tailcone.

Control switches, alarms, and emergency shutdown means areprovided on the cockpit overhead panel.

The normal APU indications and caution/warning messages arepresented on the EICAS.

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APU INSTALLATION

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CONTROL SYSTEMThe APU control systems include the electrical, fuel, ignition,lubrication, and pneumatic systems.

On APU Model T-62T-40C14, the electrical control system consists ofthe Full-Authority Digital Electronic Control (FADEC). The controlsystem incorporates the APU starting system, control logic, and failureindication. Electric accessories provide FADEC inputs and executeoutput commands.

Electrical power for the APU control is fed from two bus bars. One ofthese buses is supplied by the APU starter-generator itself, and theother is supplied by the airplane electrical system. This arrangement isprovided to ensure that a loss of the airplane electrical power duringthe APU operation will not cause the APU shutdown.

The fuel system is composed of the fuel pump, fuel solenoid valves(Main, and Maximum), acceleration control, purge valve, fuel nozzles,fuel filter, and manifold. Acceleration control provides fuel inaccordance with a preprogrammed schedule. Fuel from the right wingtank is normally used to supply the APU. Alternatively, fuel from the leftwing tank may be used by means of the crossfeed valve.

NOTE: the fuel system for the Model T-62T-40C14 APU does notinclude a start fuel solenoid valve.

The ignition system provides the electrical power necessary during theAPU starting sequence. It consists of an exciter, igniter plugs, and wiring.

The APU has a self-contained lubrication system totally integrated intothe accessory gearbox. In addition to lubrication functions, the systemprovides the required oil cooling, with no need for an external heatexchanger. A thermostat, installed in the oil tank, sends a signal to theEICAS in case the oil temperature exceeds 166°C (331°F).

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The pneumatic control system consists of a modulating bleed valveand an anti-surge valve. The modulating bleed valve maintains thebleed flow below a set value, depending on the air conditioning systemrequirements and atmospheric conditions, thus maintaining the EGTwithin acceptable levels. The anti-surge valve is controlled by theFADEC, which monitors the signal from the APU bleed valve, the AirTurbine Starter (ATS) valve, and the Environmental Control System(ECS) valve.

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APU MODEL T-62T-40C14 SCHEMATIC

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APU STARTING/OPERATIONThe APU starting cycle is initiated when the APU Master knob, locatedon the APU control panel, is moved to the ON position. At this moment,an EGT valid value is showed on EICAS. When the Master switch ismomentarily set to START, DC power is applied to the starter-generator, which will drive the APU compressor up to a speed highenough to obtain sufficient airflow for combustion. On APU ModelT-62T-40C14, at this time, the airframe fuel shutoff valve is energizedto open.

On APU Model T-62T-40C14, at approximately 3% rotor speed on theground (or 0% in flight), the FADEC supplies power to the ignition unitas well as power to open the Main Fuel Solenoid Valve, allowing fuelflow to the combustion chamber. The APU continues accelerating and,when rotor speed exceeds 50%, the FADEC deenergizes the ignitionand at 70% rotor speed the FADEC commands starter disengagement.

The APU acceleration continues by the APU own means and, 7seconds after having reached 95% rotor speed, the Maximum FuelSolenoid Valve is energized and the FADEC circuits allow electricaland pneumatic power extraction through the starter-generator and thebleed valve.

If a failure in the control system occurs, associated with an APUoverspeed, the APU Model T-62T-40C14 will automatically shutdownafter the rotating parts reach 104% speed.

The APU is shut down by pressing the APU Stop Button or by settingthe Master switch to the OFF position, whenever, a stop request signalis sent to the FADEC in order to execute the APU shutdownprocedure; the FADEC overspeed protection is tested during theFADEC power-up.

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APU FAIL APU has been automaticallyshut down.

CAUTION APU OIL LO PRESS Oil pressure is below 6 psi.APU OIL HI TEMP Oil temperature is above

166°C (331°F).

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APU CONTROL PANEL1 - APU MASTER KNOB

OFF - Deenergizes the FADEC, closes the APU fuel shutoff valve,turns off the APU indications and alarms whenever APURPM is below 10%, and commands the APU shutdown.

ON - Energizes the FADEC, commands the fuel shutoff valve toopen, enables indication and alarms on the EICAS andallows the APU to keep running after starting.

START (momentary position) - Initiates start cycle.

2 - APU STOP BUTTON− Shuts the APU down.

NOTE: APU EICAS indications remain operational.

3 - APU FUEL SHUTOFF BUTTON (guarded)− Cuts off fuel to the APU.− A striped bar illuminates inside the button to indicate that it is

pressed.

APU CONTROL PANEL

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EICAS INDICATIONS

1- APU RPM INDICATION

− Ranges from 0 to 120% speed.− Green from 96 to 104%.− Amber and boxed from 0 to 95% and

from 105 to 110%.− Red and boxed above 110%.

2- APU EGT INDICATION

− NORMAL OPERATION− Ranges from −54 to 927°C.− Green from 0 to 680°C.− Amber and boxed from 681 to 717°C.− Red and boxed above 717°C.

− START SEQUENCE− Ranges from −54 to 927°C.− Green from 0 to 838°C.− Amber and boxed from 839 to 884°C.− Red and boxed above 884°C.

NOTE: After APU shutdown, the RPM and EGT indications arereplaced by APU OFF inscription, provided the APU MasterKnob is set to OFF position and APU speed is below 10%.

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SECTION 2-04

CREW AWARENESSTABLE OF CONTENTS

Block Page

Index ................................................................................. 2-04-00 ..01

General .............................................................................. 2-04-05 ..01Avionics Integration ........................................................ 2-04-05 ..01Displays .......................................................................... 2-04-05 ..06EICAS Messages ........................................................... 2-04-05 ..18Controls and Indicators................................................... 2-04-05 ..20Built-in Test..................................................................... 2-04-05 ..29

Visual Warnings ................................................................. 2-04-10 ..01Warning Lights ............................................................... 2-04-10 ..01EICAS Messages ........................................................... 2-04-10 ..03EICAS Message Dictionary ............................................ 2-04-10 ..04Displays Indications........................................................ 2-04-10 ..11Controls and Indicators................................................... 2-04-10 ..12

Aural Warnings .................................................................. 2-04-15 ..01Aural Warning Unit ......................................................... 2-04-15 ..01EICAS Message ............................................................. 2-04-15 ..04

Takeoff Configuration Warning .......................................... 2-04-20 ..01EICAS Message ............................................................. 2-04-20 ..01Controls and Indicators................................................... 2-04-20 ..02

Stall Protection System ...................................................... 2-04-25 ..01General........................................................................... 2-04-25 ..01EICAS Messages ........................................................... 2-04-25 ..04Controls and Indicators................................................... 2-04-25 ..06

Enhanced Ground Proximity Warning System................... 2-04-30 ..01Modes and Messages .................................................... 2-04-30 ..04EGPWS Additional Features .......................................... 2-04-30 ..24Warning Priorities ........................................................... 2-04-30 ..32EICAS Messages ........................................................... 2-04-30 ..33Controls and Indicators................................................... 2-04-30 ..34Steep Approach Operation ............................................. 2-04-30 ..39

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Windshear Detection and Escape Guidance System ........2-04-35.. 01Windshear General Information......................................2-04-35.. 01Windshear Detection ......................................................2-04-35.. 04Windshear Escape Guidance Mode ...............................2-04-35.. 06EICAS Message .............................................................2-04-35.. 10Controls and Indicators...................................................2-04-35.. 10

Traffic and Collision Avoidance System .............................2-04-40.. 01General ...........................................................................2-04-40.. 01System Description.........................................................2-04-40.. 01TCAS Voice Messages...................................................2-04-40.. 08Controls and Indicators...................................................2-04-40.. 10TCAS Test ......................................................................2-04-40.. 14

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GENERALThe EMB-135 BJ is provided with a variety of visual, aural, andsensitive warnings to notify crew regarding systems status,malfunctions, and abnormal airplane configurations.

Alarm lights provide indication whether there is an abnormal situation.Some systems also provide indicating lights, for system statusindication.

An Engine Indication and Crew Alerting System (EICAS) provides theflight crew with a three-level alerting and indications messages system:warning, caution and advisory. A fourth level is provided exclusively formaintenance purposes. Besides the five displays on the main panel,two back up displays are provided through the RMUs (RadioManagement Unit). Some of the more critical messages also generatean aural warning.

Sensitive warning is available through the Stall Protection System(SPS), which shakes the control column, if an imminent stall isdetected.

To aid in navigation and approach procedures, the airplane is alsoprovided with a Enhanced Ground Proximity Warning System(EGPWS), a Traffic and Collision Avoidance System (TCAS), and aWindshear Detection and Escape Guidance System.

AVIONICS INTEGRATION

The EMB-135 BJ is equipped with a variety of highly integratedcomputers and displays, so as to reduce pilots workload whileproviding high reliability and redundancy. This feature is achieved byproviding different paths to each type of data, thus minimizing thepossibility of losing information due to failure in one computer.

The system is composed of:− Two Integrated Computers (IC-600);− Two Integrated Computer Configuration Modules (IM-600);− Two Data Acquisition Units (DAU);− One Central Maintenance Computer (CMC);− One Horizontal Stabilizer Control Unit (HSCU);− Two Primary Flight Displays (PFD), two Multi-Function Display (MFD)

and one Engine Indications and Crew Alerting System (EICAS)display;

− Two Radio Management Units (RMU);

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− One Tuning Backup Control Head;− One Integrated Standby Instruments System (ISIS);− Two Integrated Navigation Computers;− Two Integrated Communication Computers;− Three Digital Audio Panels (DAP);− Two Inertial Reference System (IRS);− Two Air Data Computers (ADC);− One Enhanced Ground Proximity Warning System (EGPWS);− One Aural Warning Unit (AWU);− One Cockpit Voice Recorder (CVR);− One Flight Data Recorder System (FDRS);− One or two Flight Management Systems (FMS);− One Traffic and Collision Avoidance System (TCAS);− One Radar System;− One Stall Protection System (SPS).

The primary components of such integration are the IC-600 units,which exchange information with all the other components, eitherdirectly or through auxiliary computers. The IC-600s are responsiblefor the interface among the many airplane systems, besides managinginformation presented on the displays. Each IC-600 computes thereceived data and sends the appropriate information to the displays.

The DAUs are the central data collection points for the EICAS. DAU 1is dedicated to collect data from the forward airplane systems and leftengine. DAU 2 collects data from the aft airplane systems and rightengine. Engine data is sent to the DAUs through the FADECs anddirectly from the engine sensors.The discrete signals collected by the DAUs are converted into digitalsignals and sent to the Integrated Computer (IC-600). In the IC 600there is a symbol generator which provides images to Display Units.Each DAU is a dual (A and B) channel unit. Channels B on both DAUsare kept as a standby source, which must be manually selected,through the DAU reversionary button in case of a channel A DAUfailure. Both IC-600s use channel A of on-side DAU as the primarysource of information.

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Normally IC 600 # 1 provides images to PFD 1, MFD 1 and EICASdisplay, while IC 600 # 2 provides images to MFD 2 and PFD 2.Both computers interact with each other and send outputs to the AuralWarning unit to generate a tone that indicates a caution or warningmessage whether there is an abnormal situation.

If IC 600 # 1 fails, the RMU # 1 displays the first engine pageautomatically. After IC 600 # 1 failure, IC 600 # 2 will control the fivedisplays by setting the Symbol Generator ("SG") button on the leftreversionary panel. In this case the RMU # 1 goes back to the normalmode.

If IC 600 # 2 fails, IC 600 # 1 will control the five displays after the "SG"button is set on the right reversionary panel. The RMU # 1 is notoperated automatically.

If both ICs fail, the RMU # 1 displays automatically the first enginebackup page.

Usually, airplane configuration options are set on the IC-600 throughstraps. If the number of installed options exceeds the maximumadjustable through the IC-600 wiring, a configuration module (IM-600)is installed. It stores information for several airplane configurations.A caution message CHK IC CONFIG appears in case of discrepancybetween the following data: aircraft id, engine type, Long Rangeconfiguration or English/Metric units. CONFIG MISMATCH message isstill active in case of discrepancy of the other parameters that do nottrigger the CHK IC CONFIG message.If a IM-600 failure occurs, the IC-600 will use the last data read fromthat source (when it was still working), and an advisory IC CONFIGFAIL message will appear.

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DISPLAYS

Five Cathode Ray Tube (CRT) displays are provided to presentinformation to the flight crew, as follows:− Two Primary Flight Displays (PFD) on the pilot and copilot panel.− Two Multi-Function Displays (MFD) on the pilot and copilot panel.− One EICAS display on the center panel.

In addition, the Radio Management Unit (RMU) displays on the mainpanel may be used as a back-up for the PFDs, MFDs and EICASdisplays.

The displays themselves are identical and interchangeable. Thecontrol panel installed just below each display, except for the RMUs,allows controlling some of the associated display features.

In case of failure of one display, its information may be presented inone of the remaining operative displays.

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PRIMARY FLIGHT DISPLAY (PFD)

The PFD is the primary pilots instrument. It presents the informationformerly presented in a variety of instruments such as airspeedindicator, altitude indicator, ADI, HSI, vertical speed indicator. The PFDfurther provides radio aids, autopilot, flight director, yaw damper andradio altitude information. For further information on these parameters,refer to Sections 2-17 − Flight Instruments, 2-18 − Navigation andCommunication, and 2-19 − Autopilot.

The PFD is divided into sections, each one presenting one group ofinformation.

The PFD bezel incorporates an inclinometer, buttons and a knob forbarometric settings.

In case of a display failure, information may be presented on the MFDby appropriately setting the MFD selector knob on the reversionarypanel.

The RMU is also able to present PFD information (refer to Section 2-18 - Navigation and Communication for further details about thisfeature).

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NOTE: Number inside boxes refer to Operations Manual Section whereinformation concerning the associated item can be found.

PFD DISPLAY SCHEMATIC

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MULTI FUNCTION DISPLAY (MFD)

The Multi Function Display (MFD) presents radar, TCAS, FMS, CMCand other navigation information and systems pages. There are fivesystem pages available:− Fuel: provides fuel system parameters and status.− Electrical: provides electrical system parameters and status.− Environmental and Ice Protection: provides air conditioning,

pneumatics, oxygen, and ice and rain protection systems parametersand status.

− Hydraulic and Brakes: provides hydraulic and brakes systems andstatus.

− Takeoff: provides takeoff data settings, oil level and doors status.

For further information on system pages, refer to each associatedsystem description.

The MFD may operate in three different presentation modes, besidesthe reversionary ones. The Map and Plan modes present navigationinformation. For further information on these, refer to Section2-18 - Navigation and Communication. The maintenance mode permitsaccess to maintenance messages, but is available only on the left MFDfor maintenance personnel when the airplane is on ground.

Selection of the different modes and pages may be made by using thecontrols located on the display bezel. Button functions are indicated inthe menus presented in the lower part of the display, just above eachbutton. Each button function changes, depending on which menu hasbeen selected. Menu selection is made by using the buttonsthemselves. If required, radar modes and TCAS information may beshown.

The MFD also operates as a back-up display for either PFD or EICAS,in case of such displays failure. Appropriate selections may be madethrough the reversionary panel.

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NOTE: Number inside boxes refer to Operations Manual Sectionwhere information concerning the associated item can be found.

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EICAS DISPLAY

The EICAS display presents analogic engine indications and somesystems parameters like flaps, landing gear, spoilers and trimpositions, total fuel quantity, APU and environmental information.

In the upper right corner, the EICAS display presents crew awarenessmessages:− Warning messages, red colored and always presented on the top of

the list.− Caution messages, amber colored and presented after warning

messages.− Advisory messages, cyan colored and presented after caution

messages.

For further information on engine indications presented in the upper leftcorner, refer to Section 2-10 − Powerplant. For information on EICASMessages, refer to the item Visual Warnings (Section 2-04-10).

In case of failure in the EICAS display, its information may bepresented on the MFD, by appropriately setting the MFD selector knobon the reversionary panel. The RMU is also capable of presentingsome EICAS information, should the need arise.

The EICAS bezel is provided with a knob to scroll messages if thesystem generates more messages than the display can present atonce.

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NOTE: Number inside boxes refer to Operations Manual Sectionwhere information concerning associated item can be found.

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RADIO MANAGEMENT UNIT

The Radio Management Unit (RMU) is provided for radio controllingpurposes, but may be used as a back-up for PFDs, MFDs and EICAS.The RMU display presents settings and modes for each radio (NAVand COMM), transponder, and TCAS. In case of failure of the mainpanel displays, the RMU may be selected to present navigation, engineor systems information, as well as some EICAS messages. Theinformation presentation however may change, due to the size of theRMU display. Also, some items of information may not be presented toavoid display overload. For further information on RMU features, referto Section 2-18 − Navigation and Communication.

RMU DISPLAY EXAMPLE

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NORMAL OPERATION

When the airplane is first energized, the system performs a self-test tocheck abnormal conditions in the displays.

On power up, the displays default information are the following:− PFD: presents EADI, EHSI, airspeed, altitude, radio altitude, vertical

speed scales, flight director mode, autopilot and yaw damper status.− MFD: presents takeoff page, system menu and navigation data in

Map format. This information is supplied as follows:− MFD 1: supplied by channel A of both DAUs through IC-600 # 1.− MFD 2: supplied by channel A of both DAUs through IC-600 # 2.

− EICAS: presents engine and fuel indications, crew awarenessmessages (if any), landing gear, flaps, spoilers, pressurization, APUand trims status. This information is supplied by channel A of bothDAUs through IC-600 # 1.

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FAILURE MODES

The system is developed to avoid absence of information in most ofthe failure combinations.

The failures that may affect the crew awareness system areassociated with electrical power supply or computer malfunctions. Inboth cases, the system architecture is such that only major failures willlead to loss of information presentation. Even in this condition, crew willstill have essential data available to safely continue the flight, usingstandby instruments.

ELECTRICAL SYSTEM FAILURES

Each display is supplied in such a way that in case of failure in one ormore electric buses, the remaining buses will still be supplying one ormore displays.This feature is achieved by supplying all displays with four differentbuses (two DC Buses and two Essential buses). Furthermore, eachpair of duplicated displays (PFDs, MFDs, and RMUs) are supplied bydifferent buses, one for each display.

COMPUTER FAILURES

Since both IC-600s receive data from duplicated sources, a singlefailure will not lead to loss of information addressed to the flight crew.In case of any source failure, the reversionary panel permits shiftingbetween existing sources, thus using cross side information. Thisfeature may be used only when the system is not capable of providinginformation through normal means.

DISPLAYS FAILURES

In case of any failure in the PFD or EICAS displays, the correspondinginformation may be presented in one of the remaining displays, byusing the reversionary panel. The MFD may present other displayinformation, but its data may not be presented in the remainingdisplays.

If all displays are lost, the RMU is capable of providing essential flightdata.

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DISPLAYS SUPPLYING SCHEMATIC

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EICAS MESSAGES

TYPE MESSAGE MEANINGDAU 1-2 ENG MISCOMP N1, N2, ITT engine

parameters read from bothengines are not matching.

DAU 1-2 SYS MISCOMP Systems parameters forsystem pages generationare not matching.

DAU 1-2 WRN MISCOMP Discrete signals for warningmessages generation readfrom the many systems arenot matching.

DAU 1 (2) A FAIL Associated DAU channel Ahas failed.

CAUTION IC 1 (2) OVERHEAT Associated temperature ofthe IC-600 is too high.

IC BUS FAIL A failure in theInterconnection Bus hasbeen detected.

IC 1 (2) WOW INOP ICs/Weight - On - Wheelsinterface not workingproperly.

CHECK PFD 1 (2) A miscomparison on theassociated PFD bus hasbeen detected.

CHECK IC 1 (2) SW Updating error on IC-600.CHK IC CONFIG Configuration module

mismatch (airplane model,engine type, LR version, andunits).

(Continued)

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(Continued)

TYPE MESSAGE MEANINGCONFIG MISMATCH Mismatch of any of the

configurations stored in theIM-600 modules exceptthose considered in theCHK IC CONFIG logic.

ADVISORYDAU 1 (2) B FAIL Associated DAU channel B

has failed.DAU 1 (2) REVERSION Associated DAU has been

commanded to operate withchannel B mode.

CMC FAIL CMC has failed.IC 1 (2) CONFIG FAIL A failure in the

configuration module of theIC has been detected.

DU 1 (2, 3, 4, 5) FANFAIL

Associated display fan hasfailed.

DU 1 (2, 3, 4, 5) OVHT Associated display unittemperature is too high.

IC 1 (2) FAN FAIL Associated IC fan has failed.

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PFD BEZEL

Provides controls that allow barometric settings in the PFD. For furtherinformation, refer to Section 2-17 - Flight Instruments.

MFD BEZEL

MAIN MENU

1 - SYSTEM BUTTON− Selects system menu.− If TCAS window is being displayed, it will be replaced by the previously selected system page.

2 - MFD BUTTON− Selects MFD menu.

3 - CHECKLIST BUTTON− This function is not enabled.

4 - TCAS BUTTON− Selects TCAS information to be presented on the MFD. For

further information refer to item TCAS presented in thissection.

− If TCAS is already selected, pressing the button restores thepreviously selected system page.

5 - WEATHER RADAR BUTTON− Selects weather radar information to be presented on the

MFD. For further information on weather radar, refer toSection 2-18 - Navigation and Communication.

6 - MAP PLAN BUTTON− When the radar is being displayed, enables the Map format

for radar presentation. For further information on weatherradar, refer to Section 2-18 − Navigation and Communication.

7 - MAP/PLAN RANGE KNOB− Allows setting the Map format range that is displayed on the

MFD. For further information on this feature, refer to Section2-18 − Navigation and Communication.

− Except for the SPDS menu, this knob function is available inall menus.

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SYS SUBMENU

1 - RETURN BUTTON− Returns to the main menu.

2 - TAKEOFF PAGE BUTTON− Selects the takeoff page to be presented on the MFD. For further information on this page refer to Section 2-2 – Equipment and Furnishings and Section 2-10 − Powerplant.

3 - ENVIRONMENTAL CONTROL SYSTEM AND PNEUMATIC PAGEBUTTON− Selects the environmental control system and pneumatic page to be presented on the MFD. For further information on this page refer to Sections 2-14 − Pneumatics, Air Conditioning and Pressurization and Section 2-16 − Oxygen.

4 - FUEL SYSTEM PAGE BUTTON− Selects the fuel system page to be presented on the MFD.− When fuel system page is being displayed, button function changes.− For further information on this page refer to Section 2-8 − Fuel.

5 - HYDRAULIC PAGE BUTTON− Selects the hydraulic page to be presented on the MFD. For further information on this page refer to Section 2-11− Hydraulic.

6 - ELECTRICAL SYSTEM PAGE BUTTON− Selects the electrical system page to be presented on the MFD. For further information on this page refer to Section 2-5 – Electrical.

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MFD BEZEL BUTTON MENU TREE

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MFD SUBMENU

1 - RETURN BUTTON− Returns to the main menu.

2 - REFERENCE SPEEDS BUTTON− Selects SPDS menu. For further information on this menu, refer to Section 2-17 – Flight Instruments.

3 - JOYSTICK BUTTON− NOTE: This function is available only when the FMS is installed.− Selects JSTK menu. For further information on this menu, refer to Section 2-18 – Navigation and Communication.

4 - AIRPORT AND NAVIGATION AIDS BUTTON− Provides selection and toggling of airport and navigation aids displays on the MFD. For further information on this feature, refer to Section 2-18 – Navigation and Communication.

5 - DATA BUTTON− Provides selection and toggling of waypoint identifier displays on the MFD. For further information on this feature, refer to Section 2-18 – Navigation and Communication.

6 - MAINTENANCE SELECTION BUTTON (LEFT MFD ONLY)− Presents maintenance messages on MFD.− Function is available only on the ground.

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EICAS BEZEL

Provides a knob to allow EICAS messages scrolling. For furtherinformation, refer to Visual Warnings in this Section.

REVERSIONARY PANEL

1 - MFD SELECTOR KNOBPFD - presents on the MFD the information normally presented on

the PFD. The PFD bezel button remains their normalfunction.

NORMAL - Normal MFD operation mode.EICAS - presents on the MFD the information normally presented

on the EICAS.

2 - ADC BUTTON− Changes the ADC information from the on-side ADC to the

cross-side ADC.− A striped bar illuminates inside the button to indicate that it is

pressed.

3 - IRS BUTTON− Changes the attitude and heading source from the on-side IRS

to the cross-side IRS.− A striped bar illuminates inside the button to indicate that it is

pressed.

4 - SYMBOL GENERATOR BUTTON− Changes the symbol generator from the on-side SG to the

cross-side Symbol Generator as well ADC and IRS.− A striped bar illuminates inside the button to indicate that it is

pressed.

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REVERSIONARY PANEL

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EICAS REVERSIONARY PANEL

1 - DAU REVERSIONARY BUTTON− Allows channel B of associated DAU to supply both IC-600s.− A striped bar is illuminated inside the button to indicate that it is

pressed and that channel B is the current data source.

EICAS REVERSIONARY PANEL

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PRIMARY FLIGHT DISPLAY

1 - SYMBOL GENERATOR REVERSION ANNUNCIATION− Indicates that a symbol generator reversion has been selected

on the reversionary panel.− Presented on both PFDs.− Labels: SG1 for IC-600 # 1 and SG2 for IC-600 # 2.− Color: amber


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NOTE: For further information on displays control panel, refer toSections 2-17 – Flight Instruments and 2-18 – Navigation andCommunication .

1 - TEST BUTTON− On the ground:

− When pressed, activates the IC-600s first level test.− When pressed for more than 6 seconds activates the

IC 600s second level test.− When released, normal operation of IC-600s is resumed.

− In flight: Refer to Radio Altimeter description on Section 2-17 – Flight Instruments.


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BUILT-IN TEST

There are 3 kinds of Built-In-Tests (BIT) that the IC-600 may perform:power up BIT, continuous BIT and pilot initiated BIT. All of them checkthe software and hardware integrity and operation.

POWER UP BIT

The power up BIT checks the power supply, IC-600 interfaces,memories, autopilot engagement and disengagement, and autopilotservos.

CONTINUOUS BIT

Memories and processors tests are continuously performed after thepower up BIT, as well as autopilot functions.

PILOT INITIATED BIT

A pilot initiated BIT may be commanded by pressing the TEST buttonin the displays control panel. This test may be commanded on theground only and is divided into two levels. The first level is indicated onairplane displays, which present the failure mode annunciations.The second level is activated if the TEST button is held pressed for atleast 6 seconds, and checks the IC-600 internal interfaces. The testresults are displayed on the PFD, which alternates every 10 secondsbetween internal and external test results pages.To perform the IC-600 test is necessary to press the TEST buttonlocalized at the display control panel.

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The PFD first level test indications are as follows:

− A magenta TEST is displayed in upper left center of the PFD.− Indications removed: all bugs, flight director information, all

pointers, low airspeed awareness, take-off speed bugs and digitalreadouts, VMO/MMO, and trend vectors.

− Indications forced: all comparison monitors, all marker beacons,and windshear annunciation.

− Indications presented as invalid: pitch and roll, vertical and lateraldeviations, baro correction, vertical speed set digital readout,altitude preselect, heading, distance digital readout, ground speed(or time to go or elapsed time), selected heading and course (ordesired track), Mach, airspeed, airspeed set digital readout, altitude.

− If heading is valid upon test activation, its source annunciation willremain valid (DG1 or 2 or MAG1 or 2). If heading is invalid, itssource annunciation will change to HDG1 or HDG2.

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01

PFD TEST INDICATIONS - FIRST LEVEL

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PFD TEST INDICATIONS - SECOND LEVEL

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The MFD test indications are as follows:− Indications removed: heading source, TCAS, weather patch,

drift bug, wind vector, heading select bug, flight plan data,airports, navaids, designator information.

− Indications forced: TERRAIN FAIL, EICAS CHK, WX TERRAIN,MENU INOP, HDG FAIL.

− Indications presented as invalid: heading, weather radar tilt,SAT, true airspeed, ground speed, distance and time towaypoint.

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MFD TEST INDICATIONS

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The EICAS test is commanded only from the pilot's panel, and itsindications are as follows:− Indications removed: reversion, ignition, FADEC in control, all

engine and trim bugs.− Indications forced: the crew awareness field will be filled with a

"X".− Indications presented as invalid: landing gear status, N1, N2,

ITT, fuel flow and quantity, oil pressure, temperature andquantity, vibration for LP and HP, flaps, spoilers, all cabin andAPU parameters, all trim values.

During IC-600 # 1 first level pilot initiated BIT, RMU 1 will displaythe first page of standby engine indication. The RMU 2 is notincluded in the IC-600 # 2 first level pilot initiated BIT.

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EICAS TEST INDICATIONS

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01

VISUAL WARNINGSVisual warnings are provided through lights, illuminated buttons,EICAS messages and displays indications.

WARNING LIGHTS

Some of the airplane systems are actuated by illuminated buttons.When under normal operating conditions, such buttons are notilluminated. If the pilot has commanded the button to a position thatrequires crew attention, a striped bar is illuminated inside the button.There are some exceptions such as the GPU, the ice protection wingand stab, and the APU bleed buttons, which are illuminated undernormal operating conditions.

Some systems also provide indicating lights, for system statusindication. Further details on such lights are provided in the associatedsystems description section.

Master warning and caution lights are installed on each pilotglareshield panel. Such lights blink when any warning or cautionmessage is presented on the EICAS or generated in the Aural WarningUnit (AWU). To stop blinking, pilots must press the associated light. Tofind information on illuminated buttons and any specific warning light,refer to the associated system’s description.

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EICAS MESSAGES

EICAS messages are presented in the upper right corner of the EICASdisplay. In case of a simultaneous failure in the EICAS and MFDdisplays, the RMUs are capable of presenting some messages.

EICAS MESSAGES LEVELS

There are three message levels: warning, caution, and advisory:− Warning messages are red colored and require immediate crew

action. Warning messages are always presented on the top of thelist, in the same order they are generated.

− Caution messages are amber colored and require immediate crewawareness. They follow warning in criticality level and in displaypresentation.

− Advisory messages are cyan and are dedicated to minor failures orsystem status. Advisory messages are displayed below cautionmessages.

A fourth level is provided for maintenance purposes, but it is notpresented to the flight crew, and its access can only be made on theground.

When the message is generated, it is displayed blinking at the top ofthe associated group. To stop blinking, press the associated masterbutton on the glareshield. Advisory messages will stop blinking after 5seconds.

EICAS MESSAGES PRIORITY LOGIC

If more than one message is simultaneously presented, warning willprecede caution messages, which will precede advisories. The spaceis provided for the simultaneous display of up to 15 messages. AnEND label is provided after the last message, to indicate end ofmessage listing. If more than 15 messages are being generated, aknob in the display bezel permits paging through the remainingmessages. In this case, a status line is provided in the sixteenth line, toindicate how many messages are not being currently presented andwhere they are (above or below the currently presented messages).END label and warning messages can not be scrolled out of thedisplay. Due to this characteristic, caution and advisory messages willbe scrolled in the area left blank below the warning messages. If a newmessage is generated during a scrolling, it will be automaticallydisplayed at the top of the associated group.

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01 REVISION 6

INHIBITION LOGIC

To avoid its nuisance effect upon the flight crew, inhibition logic isprovided to prevent some messages from being displayed duringtakeoff and approach/landing phases. The inhibition logic is as follows:

Takeoff Phase:

Inhibition is valid when the airplane crosses V1 –15 kt. The inhibition isdeactivated when one of the following conditions is accomplished:

− radio altitude is greater than 400 ft or;

− calibrated airspeed is less than 60 kt or;

− after 1 minute.

Approach/landing Phase:

Inhibition is valid from the point when airplane crosses 200 ft radioaltitude. The inhibition is deactivated when one of the followingconditions is accomplished:

− airplane is on the ground for 3 seconds or more;

− after 1 minute.

IC-600 RESULTS SELF-CHECK

The results of both IC-600 computations are continuously compared tocheck for any inconsistency between both sides. A dedicated amberannunciation, “CAS MSG”, is provided on the PFDs to indicatewhenever a difference between both IC-600s has been found, thusleading to possible unreliable messages.

EICAS MESSAGE DICTIONARY

The following table presents all the EICAS messages. Type columnindicates whether the message’s nature is Warning (W), Caution (C),or Advisory (A).The number in column INHIBITION indicates the following:− (1) Message is inhibited during takeoff;− (2) Message is inhibited during takeoff and approach/landing;− (3) Message is not inhibited;− (4) Message is inhibited during approach/landing;

For further information regarding each message’s logic, refer to theassociated system’s description.

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SECTION TYPEMESSAGE INHIBITION

W MAIN DOOR OPN 22-2 W SERVICE DOOR OPN 2

EQUIPMENT C BAGG ACCESS OPN 2AND C ACCESS DOORS OPN 2

FURNISHINGS C BAGGAGE DOOR OPN 2C EMERG EXIT OPN 2C FUELING DOOR OPN 2W GPWS 3W NO TAKEOFF CONFIG 4W SPS 1 (2) INOP 3C DAU 1-2 ENG MISCOMP 2C DAU 1-2 SYS MISCOMP 2C DAU 1-2 WRN MISCOMP 2C DAU AC ID MISCMP 2C AURAL WARN FAIL 2C CHECK PFD 1 (2) 2C CHK IC CONFIG 2

2-4 C CHECK IC 1 (2) SW 2CREW C DAU 1 (2) A FAIL 2

AWARENESS C GPWS INOP 3C IC 1 (2) OVERHEAT 2C IC BUS FAIL 2C IC 1 (2) WOW INOP 2C SPS ADVANCED 3C STICK PUSHER FAIL 3C WINDSHEAR INOP 3A IC 1 (2) CONFIG FAIL 2A CONFIG MISMATCH 2A CMC FAIL 2A DAU 1 (2) B FAIL 2A DAU 1 (2) REVERSION 2A DU 1 (2, 3, 4, 5) FAN FAIL 2A DU 1 (2, 3, 4, 5) OVHT 2A IC 1 (2) FAN FAIL 2

A SPS/ICE SPEEDS 1

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W BATT 1 (2) OVTEMP 3W ELEC ESS XFR FAIL 3C 115 VAC BUS OFF 2C APU CNTOR CLSD 2C APU GEN OFF BUS 2C APU GEN OVLD 2

2-5 C BATT 1 (2) OFF BUS 2ELECTRICAL C BKUP BATT OFF BUS 2

C DC BUS 1 (2) OFF 2C ELEC EMERG ABNORM 2C ESS BUS 1 (2) OFF 2C GEN 1 (2, 3, 4) OFF BUS 2C GEN 1 (2, 3, 4) OVLD 2C SHED BUS 1 (2) OFF 2A GEN 1 (2, 3, 4) BRG FAIL 2

2-6LIGHTING C EMERG LT NOT ARMD 2

W APU FIRE 3W BAGG SMOKE 2W ENG 1 (2) FIRE 3

2-7 W LAV SMOKE 2FIRE C APU EXTBTL INOP 2

PROTECTION C APU FIREDET FAIL 2C BAGG EXTBTL INOP 2C E1 (2) EXTBTLA (B) INOP 2C E1 (2) FIREDET FAIL 2W FUEL 1 (2) LO LEVEL 2W CHECK ACFT LOAD 2W NO TAKEOFF CONFIG 4W FUEL XFER CRITICAL 2C APU FUEL LO PRESS 2C APU FUEL SOV INOP 2

2-8 C E1 (2) FUEL LO PRESS 2FUEL C E1 (2) FUEL SOV INOP 2

C FUEL TANK LO TEMP 2C FUEL TK VENT OPEN 2C FUEL XFEED FAIL 2C FUEL XFEED MISCMD 2C FUEL XFER CHECK 2

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C FUEL XFER 1(2) INOP 2C FUEL XFER OVFLOW 2

2-8 C FUSELAGE FUEL IMB 2FUEL C WING FUEL IMB 2

(Cont.) A APU FUEL SOV CLSD 2A E 1(2) FUELSOV CLSD 2A FUEL XFEED OPEN 2

2-9 C APU FAIL 2APU C APU OIL HI TEMP 2

C APU OIL LO PRESS 2W ATTCS FAIL 4W E1 (2) ATTCS NO MRGN 2W E1 (2) OIL LOW PRESS 2W E1 (2) LOW N1 2W ENG 1-2 OUT 1C E1 (2) EXCEEDANCE 2C E1 (2) FPMU NO DISP 2C E1 (2) ATS SOV OPN 2C E1 (2) CTL FAIL 2C E1 (2) FUEL LO TEMP 2C ENG NO TO DATA 2

2-10 C ENG REF A/I DISAG 2POWERPLANT C ENG1 (2) REV DISAGREE 2

C E1 (2) NO DISP 2C ENG 1 (2) OUT 1C FADEC ID NO DISP 2C ENG 1 (2) REV FAIL 2C ENG 1 (2) TLA FAIL 2A CHECK A1P PERF 2A CHECK A1E PERF 2A E1 (2) SHORT DISP 2A E1 (2) FUEL IMP BYP 2A E1 (2) OIL IMP BYP 2A E1 (2) IDL STP FAIL 2

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C HYD SYS 1 (2) FAIL 2C HYD SYS 1 (2) OVHT 2

2-11 A E1 (2) HYD PUMP FAIL 2HYDRAULIC A E1 (2) HYDSOV CLSD 2

A HYD PUMP SELEC OF 2A HYD1 (2) LO QTY 2W LG/LEVER DISAGREE 3C BRAKE OVERHEAT 3

2-12 C BRK INBD INOP 1LANDING C BRK OUTBD INOP 1

GEAR AND C EMRG BRK LO PRES 1BRAKES C LG AIR/GND FAIL 3

C STEER INOP 2C NLG UP/DOOR OPN 2C BRAKE DEGRADED 1W PTRIM MAIN INOP 2W PTRIM BACKUP INOP 2C AIL SYS 1 (2) INOP 2C FLAP FAIL 2C PTRIM CPT SW FAIL 3

2-13 C PTRIM F/O SW FAIL 3FLIGHT C PTRIM BKP SW FAIL 3

CONTROLS C RUDDER OVERBOOST 2C RUDDER SYS 1 (2) INOP 2C RUD HDOV PROT FAIL 2C SPBK LVR DISAGREE 2C SPOILER FAIL 2A FLAP LOW SPEED 2W BLD 1 (2) LEAK 2W BLD APU LEAK 2

2-14 W BLD 1 (2) OVTEMP 2PNEUMATICS, C APU BLD VLV FAIL 2

AIR C BLD 1 (2) LOW TEMP 2CONDITIONING C BLD 1 (2) VLV FAIL 2

AND C CROSS BLD FAIL 2PRESSURIZATION C CROSS BLD SW OFF 2

C ELEKBAY OVTEMP 2C HS VLV 1 (2) FAIL 2

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C PACK 1 (2) OVHT 22-14 C PACK 1 (2) OVLD 2

PNEUMATICS, C PACK 1 (2) VLV FAIL 2AIR CONDITIONING C PRESN AUTO FAIL 2

AND C RAM AIR VLV FAIL 2PRESSURIZATION A BLD 1 (2) VLV CLSD 2

(Cont.) A CROSS BLD OPEN 2A PACK 1 (2) VLV CLSD 2W ICE COND-A/I INOP 2C A/ICE SWITCH OFF 2C A/ICE LOW CAPACITY 3C AOA 1 (2) HEAT INOP 2C CLR ICE 1 (2) 2C CLR/I INOP 1 (2) 2C E1 (2) A/ICE FAIL 2

2-15 C ICE DET1 (2) FAIL 2ICE AND RAIN C ICE DETECTORS FAIL 2PROTECTION C NO ICE-A/ICE ON 2

C PITOT 1 (2, 3) INOP 2C STAB A/ICE FAIL 2C TAT 1 (2) HEAT INOP 2C WG A/ICE FAIL 3C W/S 1 (2) HEAT FAIL 2A ICE CONDITION 2A ENG A/ICE OVERPRES 3

2-16 C CREW OXY LO PRESS 2OXYGEN C PAX OXY LO PRESS 2

2-17 C DFDR FAIL 2FLIGHT A FDAU FAIL 2

INSTRUMENTS A RAD ALT 1 (2) FAIL 1A RAD ALT FAIL 1

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C IRS 1 (2) ALN FAULT 2C IRS 1 (2) FAIL 2

2-18 C IRS 1 (2) OVERHEAT 2NAVIGATION A IRS 1 (2) ALN 2

AND A IRS 1 (2) ATT MODE 2COMMUNICATION A IRS 1 (2) EXC MOTION 2

A IRS 1 (2) NO MAG HDG 4A IRS 1 (2) NO PPOS 1A IRS 1 (2) ON BATT 2W AUTOPILOT FAIL 2C AUTO TRIM FAIL 2

2-19 C AP ELEV MISTRIM 2AUTOPILOT C AP AIL MISTRIM 2

C LATERAL MODE OFF 3C VERTICAL MODE OFF 3C YAW DAMPER FAIL 2

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DISPLAYS INDICATIONS

Many of the airplane’s parameters are indicated on one of the displays,in analogic or digital format.

ANALOGIC INDICATIONS

Analogic indications are provided as pointers moving over a scale,which may be graduated or not. In both cases, if the pointer indicates avalue out of the normal range for that parameter, both pointer andscale become amber or red, if the parameter goes deeply into the outof range area. Pointers are removed if the parameter signal becomesinvalid. For some parameters, scale may also be removed in thiscondition. Scale and pointer are not presented for some parameters,when they are not required, as for EADI chevrons, V1, VR, V2 speedbugs, trend vectors, etc.

DIGITAL INDICATIONS

Digital indications are provided as green characters for normal values.If the associated parameter goes outside its normal range, digitsbecome amber, with an amber box surrounding them. Both digits andbox become red if the parameter goes deeply into the out of rangearea. If the parameter signal becomes invalid, digits are replaced byamber dashes, without boxes.

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GLARESHIELD PANEL

1 - MASTER WARNING BUTTON

− Acknowledges the warning messages and stops the associatedblinking when pressed.

− A red light blinks inside the button when a new warningmessage is displayed on the EICAS.

2 - MASTER CAUTION BUTTON

− Acknowledges the caution messages and stops the associatedblinking when pressed.

− An amber light blinks inside the button when a new cautionmessage is displayed on the EICAS.

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GLARESHIELD PANEL

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EICAS BEZEL

1 - MESSAGE SCROLLING KNOB− To be used when displayed EICAS messages can not be

presented at once.− By rotating the knob clockwise, advances through EICAS

messages. Rotated counterclockwise moves backward throughEICAS messages.

EICAS BEZEL

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1 - EICAS CHECK SUM FAIL COMPARISON MONITOR DISPLAY− Color: amber.− Label: CAS MSG.− Displayed when the number of active EICAS messages in each

IC-600 is found to be different.


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EICAS DISPLAY

EICAS MESSAGES EXAMPLE

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RMU DISPLAY

RMU MESSAGES EXAMPLE

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AURAL WARNINGSThere are two kinds of aural warnings: voice messages and tones.

Voice messages are normally associated with warning messages onEICAS or other warning systems. They are generated whenever apotentially dangerous condition exists, as determined by the EGPWS,TCAS and windshear detection system. There are some voicemessages that can be cancelled, but others can only be cancelledwhen the cause that triggered them has been eliminated.

Tones have different forms and indicate some notable airplane events,sometimes in unison with voice messages.

AURAL WARNING UNITIn order to generate messages and tones, the Aural Warning Unit(AWU) receives signals from the following airplane systems:

− TCAS− Windshear detection system− EGPWS− IC-600− Fire detection system− Stall protection system− Trims− Flaps− Brakes− Spoilers− Radio altimeter− Autopilot− Landing gear− ADC− Pressurization− SELCAL

The AWU sends the appropriate audio signal to an audio digitalsystem, which routes the messages to the appropriate speakers.

AWU POWER SOURCE

The AWU is supplied by one DC bus and one Essential DC bus, and isprovided with two channels, A and B. Channel B is kept as a backupfor channel A and is automatically activated if channel A fails.

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AWU POWER-UP TEST

An AWU power-up test is performed and generates aural warnings forone or both channels operating normally. If both channels have failed,the caution message AURAL WARN FAIL is displayed on EICAS.

AURAL WARNINGS LEVELS

The aural warnings are classified into four levels, presented below in adecreasing level order:

− Emergency - Associated with situations that may be hazardous.AWU generates a master warning tone (triple chime) before thewarning and voice message may be generated. In any case, theaural warning is repeated every second until deactivated through themaster warning button or until the condition that generated thewarning has been eliminated.

− Abnormal - Associated with malfunctions or failures. AWUgenerates a master caution tone (single chime) every five seconds,until it is removed, canceled or replaced by a higher priority auralwarning. Voice messages are generated after each tone.

− Advisory - Associated with minor malfunctions or failures that lead toloss of redundancy or degradation of the affected system’s performance.

− Information - A remarkable event has occurred.

AURAL WARNINGS ANNUNCIATION PRIORITY

When multiple aural warnings are active, aural warnings among thehighest level alert groups shall be sounded first in order and repeated.Once all alerts in the higher group are cancelled or removed, then thesecond tier group alerts are sounded and repeated.

An alert in process shall be immediately interrupted when an alert of ahigher priority needs to be generated.

EXCEPTIONS TO AURAL WARNINGS PRIORITY

When an internal voice message is being annunciated, it shall becompleted before another alert, even of a higher priority, isannunciated. This does not apply to internally generated tones whichshall be interrupted within 1 second.

If an emergency arises together with a warning that generatescontinuous sounds, such as a fire or stall, the sound volume is reducedto avoid misunderstanding of the remaining messages, although beingloud enough to still warn pilots.

The master warning tone is inhibited when any other emergency alert(internal or external) is occurring at the same time.

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LEVELASSOCIATED

CONDITION/EICASMESSAGE

PRIORITY TONE VOICEMESSAGE

CANCEL

Stall condition. 1 Clacker None NOWindshear condition (1). 2 None WINDSHEAR NOGround proximity condition(1).

3 (1) (1) NO

Traffic proximity condition(1).

4 None (1) NO(2)

Fire in engine or APU.ENG 1 (2) FIRE ,APU FIRE.

5 Bell None YES

Airspeed above VMO. 6 Attenson3

HIGHSPEED

NO

EMERGENCY Landing gear not lockeddown for landing.

7 Attenson3

LANDINGGEAR

NO

Cabin altitude above10000 ft.

8 Attenson3

CABIN YES

Associated with takeoffconfiguration warning.

9 Attenson3

TAKEOFFplus one ofthe following:

- BRAKES

- FLAPS

- FUEL

- SPOILER

- TRIM

NO

Associated withemergency failures.

10 Attenson3

None NO

Associated with glideslope deviation.

None None GLIDESLOPE

YES

ABNORMAL Traffic proximity condition. None None TRAFFIC YESAssociated with abnormalfailures.

None MasterCaution

Tone

None YES

NOTE: 1) Messages are generated outside the AWU. For furtherdetails, refer to the associated system description.

2) TCAS resolution advisory warning can not be canceled.3) Applicable to airplanes equipped with CMU.

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LEVELASSOCIATED

CONDITION/EICASMESSAGE

PRIORITY TONE VOICEMESSAGE

CANCEL

Autopilot disengagementduring approach.

None None AUTOPILOT NO

ADVISORYAssociated with decisionheight crossing.

None None MINIMUM NO

Airplane is crossing orhas reached apreselected altitude.

None Three2900 Hztones

None NO

Power up test detected afailure in one channel ofAWU.

Notapplicable

None AURALUNIT ONECHANNEL

NO

Associated with incorrectcommand of pitch trimmain or backup channelswitches.

None Singlechime

TRIM NO

INFORMATIONAssociated with SELCALcallings.

None None SELCAL NO

Both AWU channels areoperating normally onpower up test.

None None AURALUNIT OK

NO

Takeoff configuration testsuccessful.

None None TAKEOFFOK

NO

Power 1 or 2 fail. None None AURALUNIT ONEPOWER

Notapplicable

When CMU receives anew message.

None None INCOMINGCALL

(3)

NO

NOTE: 1) Messages are generated outside the AWU. For furtherdetails, refer to the associated system description.

2) TCAS resolution advisory warning can not be canceled.3) Applicable to airplanes equipped with CMU.

EICAS MESSAGE


CAUTIONAURAL WARN FAIL Both AWU channels are

inoperative.

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TAKEOFF CONFIGURATION WARNINGA dedicated warning to indicate that airplane configuration is unsuitablefor takeoff is provided by the aural warning system. Such warning isactivated whenever the airplane is on the ground, thrust is applied andat least one of the following conditions is met:

− Parking brakes are applied.− Flaps are not in takeoff position.− Fuselage Tank Transfer System Knob is not in off position.− Any spoiler panel is deployed.− Pitch trim is out of the green range.

More than one warning may be generated, if more than one conditionis met.

TEST BUTTON

A test button is provided to allow checking the takeoff configurationwarning integrity, by simulating power levers advanced. A voicemessage is generated after successful tests. Unsuccessful tests willgenerate an EICAS message and a voice message associated with theout-of-configuration item.

EICAS MESSAGE


WARNING NO TAKEOFF CONFIGAirplane is not in takeoffconfiguration.

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1 - TAKEOFF CONFIGURATION CHECK BUTTON

− Allows checking the takeoff configuration warning.

TAKEOFF CONFIGURATION CHECK BUTTON

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STALL PROTECTION SYSTEM

GENERAL

To help detect imminent stalls and to avoid stalling the airplane, theEMB-135 BJ is provided with a Stall Protection System (SPS). TheSPS is composed of one computer box with two independentchannels, the SPS panel, two Angle of Attack (AOA) sensors, two stickshaker actuators, and one stick pusher actuator. The system providessensitive, visual and aural indications of an impending stall. To avoidspurious actuation, the SPS receives signals from many airplanesystems, thus correcting its set point according to flaps and landinggear position, icing and windshear conditions and Mach number.

INTERFACES

Each channel receives data from the following on-side airplanesystems: IRS, ADC, flaps, landing gear, air/ground, windsheardetection, ice detection and radio altimeter. Each Stall ProtectionComputer (SPC) channel receives information from its associated AOAsensor and sends it to the opposite channel in order to compensateside slip influence on angle of attack measurements. A locked AOAsensor signal is not considered in stall calculations and in this case thechannel will be deactivated. If a stall condition is imminent, the systemfirst actuates the stick shaker and disengages the autopilot. If nocorrective action is taken and the airplane is on the verge of entering astall, the stick pusher is actuated, which pitches the nose down.Simultaneously, a clacker is generated in the aural warning system. Abug in the airspeed scale on the PFD indicates the stall speed for theassociated condition and a pitch limit indicator is presented on EADI toindicate the current margin to the stick shaker angle. When theairplane reaches 0.5 g, the stick pusher is inhibited, stopping itsactuation over the control column. A quick disconnect button isprovided in the control wheel to permit pilots to cut the system if theneed arises. To disconnect the system in case of failure, the SPSpanel provides one cutout button for each channel. An EICASmessage is presented to indicate that the system has failed or iscutout.

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SYSTEM INHIBITION

The stick pusher does not actuate in the following conditions:− On the ground (except during test).− Below 0.5 g.− If the quick disconnect button is pressed (except for JAA certification).− Below 200 ft AGL. If radio altimeter has failed, this condition reverts

to a 10-second delay after takeoff.− If any cutout buttons are released.− Above 200 KIAS.− If at least one channel is inoperative.

SYSTEM TEST

A test button is provided to test the system on the ground. The systemoperates normally if not tested. Test button remains illuminated if thesystem has not been tested or after unsuccessful tests. It is not possibleto test the system in flight. This inhibition is valid for 30 seconds afterlanding, above 70 KIAS or with landing gear not downlocked.

NOTE: Test button remains illuminated if quick disconnect button ispressed during test.

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STALL PROTECTION SYSTEM SCHEMATIC

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EICAS MESSAGES

TYPE MESSAGE MEANINGSPS 1(2) INOP Associated SPS computer channel

has failed or AOA vane failed.WARNING SPS 1-2 INOP Both SPS computer channels

have failed or both AOA vaneshave failed or stick pusher hasfailed or is cutout.

CAUTION

SPS ADVANCED Stick shaker and pusheractuation is set to higher speedsdue to:− Flap signal disagreement.− Failure in at least one SPS

channel.− IRS or ADC parameters

disagree.− Air/Ground signs disagree.− Landing gear down and locked

indications disagree.STICK PUSHER FAIL Stick pusher actuator has been

commanded but has not moved.

ADVISORYSPS/ICE SPEEDS SPS actuation angle is advanced

for flaps 9° and 22°.

NOTE: Advisory SPS/ICE SPEEDS messages are inhibited for the first5 minutes after takeoff.

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CONTROLS AND INDICATIORS

STALL PROTECTION SYSTEM PANEL

1 - CUTOUT BUTTON (guarded)

− Cuts out the associated channel.− A striped bar illuminates inside the button to indicate that it is in

the cutout position.

2 - TEST BUTTON

− Starts the test sequence, as follows:− Button illuminates.− Both stick shakers actuate.− Pusher actuates.− Button illumination extinguishes.NOTE: - Test sequence is completed within a maximum of 5

seconds.- The TEST button must be released at the first sign of

stick shaker actuation.

− Button is kept illuminated after an unsuccessful test or if thesystem has not been tested.

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STALL PROTECTION SYSTEM PANEL

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PFD INDICATIONS

1 - PITCH LIMIT INDICATOR

− Displayed on the EADI parallel to the airplane symbol.− Indicates the remaining margin left for the stick shaker angle of

attack set point.− Indication is presented whenever the margin reaches 10°.− Color:

− green for margin from 10° up to 5°.− amber for margin between 5° and 2°.− red for margin below 2°.

2 - LOW AIRSPEED AWARENESS

− Displayed in the airspeed scale when airspeed is near stallspeed for the current configuration.

− For further details on Low Airspeed Awareness, refer to Section2-17–Flight Instruments.

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ENHANCED GROUND PROXIMITY WARNINGSYSTEMThe purpose of the Enhanced Ground Proximity Warning System(EGPWS) is to avoid accidents caused by Controlled Flight Into Terrain(CFIT) and also severe windshear.

The EGPWS incorporates functions like Terrain Clearance Floor,Terrain Look Ahead Alerting and Terrain Awareness Display. Thesefunctions use airplane geographic position, airplane altitude and aninternal terrain database to predict potential conflicts between theairplane's flight path and terrain, and to provide graphic displays of theconflicting terrain.

NOTE: Airplanes equipped with EGPWS version 216 incorporatesadditional features like Peaks Mode, Runway Field ClearanceFloor, Obstacle Alerting and Geometric Altitude.

The EGPWS is a useful navigation aids when flying at low altitude,generally within 2500 ft above terrain. It provides voice messages,EICAS message and PFD indication to alert the flight crew, so thatthey may take appropriate action.

The EGPWS interfaces with the followings systems and equipment:

− Radio Altimeter - The radio altimeter provides altitude above ground,how fast the altitude decreases as a result of airplane sinkage orground profile change and the validity signal.

− IC-600s - The IC-600s provide glideslope deviation, localizerdeviation, selected decision height, selected course, packed discreteand selected terrain range.

− ADCs - The ADCs provide uncorrected barometric altitude, correctedbarometric altitude, computed airspeed, true airspeed, barometricaltitude rate and static air temperature.

− IRS - The IRS provide magnetic heading, pitch and roll angle,longitudinal and normal acceleration.

− FMS - The FMS provides latitude, longitude, ground speed, truetracking, true heading and NAV mode. The same is applicable whenthe airplane is equipped with dual FMS.

− GPS - The GPS provides latitude, longitude and altitude.− Landing gear - The landing gear provides a discrete signal that

indicates gear down/locked condition.− Flap - The Flap Control Unit provides one discrete signal that

indicates whether or not flaps are in landing position.

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− AWU - The AWU receives the aural messages to be enunciated. Italso provides a discrete signal to indicate that the glideslope advisoryalert may be canceled without any restriction.

− Terrain Inhibit Switch - It is used in approach mode, in airports notcovered by an EGPWS database, assuring protection againstunwanted terrain alerts.

Some modes may have their associated envelopes shifted, so as tosuit particular airport requirements or to avoid nuisance warningsunder some flight situations. This feature is achieved either withcalculations or data provided by the FMS, if installed.

The EGPWS provides alerts associated with the following flightconditions:

Mode 1 - Excessive descent rate.Mode 2 - Excessive closure rate to terrain.Mode 3 - Altitude loss after takeoff.Mode 4 - Insufficient terrain clearance.Mode 5 - Excessive deviation below glideslope beam.Mode 6 - Callouts.Mode 7 - Windshear (refer to Section 2-04-35).Terrain awareness alerting and warning.Terrain clearance floor.

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EGPWS SCHEMATIC

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MODES AND MESSAGES

MODE 1 - EXCESSIVE DESCENT RATE

Mode 1 provides alerts and warnings when the airplane has attainedan excessive descent rate in respect to altitude above ground level(AGL) during the descent and approach phases of flight.

This mode has outer (sink rate) and inner (pull up) alert/warningboundaries:

Minimum Terrain Clearance (MTC) for “SINK RATE” messagetriggering:− Minimum: 30 ft at 1000 ft/min of descent Altitude Rate.− Maximum: 2450 ft at 5007 ft/min or greater of descent Altitude

Rate.

Minimum Terrain Clearance (MTC) for “PULL UP” message triggering:− Minimum: 30 ft at 1710 ft/min of descent Altitude Rate.− Maximum: 2450 ft at 7125 ft/min or greater of descent Altitude

Rate.

Penetration of the outer (sinkrate) boundary will result in:− Aural message “SINK RATE”. The message will be repeated as

long as the penetration increases; and− Amber "GND PROX" indication on the PFD.

Penetration of the inner (pull up) boundary trigger and repeat thefollowing messages until the condition is cleared:− Aural message “PULL UP” and red "PULL UP" indication on the

PFD.

If a valid ILS Glideslope front course signal is received and the airplaneis above the glideslope centerline, the sinkrate boundary is adjusted toprevent unwanted alerts when the airplane is safely capturing theglideslope.

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EGPWS MODE 1 SCHEMATIC

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MODE 2 - EXCESSIVE CLOSURE RATE TO TERRAIN

Mode 2 provides alerts and warnings based on airspeed, airplanegear/flap configuration, radio altitude, and excessive closure rate toterrain. Mode 2 exists in two forms: 2A and 2B.

MODE 2A

Mode 2A is selected when the flaps are not in landing configurationand the airplane is not on the glide slope beam.

Minimum Terrain Clearance (MTC) for “TERRAIN TERRAIN”message triggering:− Minimum: 30 ft at 2038 ft/min of Closure Rate.− Maximum:

− 1650 ft at 5733 ft/min or greater of Closure Rate, for an airspeed equal or below 220 KIAS.− 2450 ft at 9800 ft/min or greater of Closure Rate for an airspeed equal or above 310 KIAS.

If the airplane penetrates the Mode 2A envelope, the situation resultsin:− Aural message “TERRAIN, TERRAIN” ; and− Amber "GND PROX" indication on the PFD.

If the airplane continues to penetrate the envelope, the aural messageswitches to messages described below, until the condition is cleared:− Aural message “PULL UP” and red "PULL UP" indication on the

PFD.

The visual and aural messages will remain on until the airplane hasgained 300 ft of barometric altitude.

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EGPWS MODE 2A SCHEMATIC

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MODE 2B

Mode 2B is selected when the flaps are in landing configuration orwhen making an ILS approach with glide slope and localizer deviationsbelow 2 dots.

Minimum Terrain Clearance (MTC) for “TERRAIN TERRAIN”message triggering:− Minimum: 30 ft at 2038 ft/min of closure rate.− Maximum:

− 789 ft at 3000 ft/min or greater of closure rate. This steady value can also vary from 200 ft up to 600 ft for flaps set to landing configuration.

If the airplane penetrates the Mode 2B envelope with both gear andflaps in the landing configuration, the message “TERRAIN” is sounded.

If the airplane penetrates the mode 2B envelope with either the landinggear UP or flaps not in landing configuration will result in:− Aural message “TERRAIN, TERRAIN” ; and− Amber "GND PROX" indication on the PFD.

If the airplane continues to penetrate the envelope, the aural messageswitches to messages described below, until the condition is cleared:− Aural message “PULL UP” and red "PULL UP" indication on the

PFD.

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EGPWS MODE 2B SCHEMATIC

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MODE 3 - ALTITUDE LOSS AFTER TAKEOFF

Mode 3 provides alerts and warnings for a significant altitude loss aftertakeoff with landing gear UP or flaps in other than landingconfiguration. The amount of altitude loss required to trigger thewarning depends on the height of the airplane above the terrain.

Minimum Terrain Clearance (MTC) for “DON'T SINK, DON'T SINK”message triggering:− Minimum: 30 ft at 5 ft of altitude loss.− Maximum: 1500 ft at 143 ft or greater of altitude loss.

Significant altitude loss after takeoff or during a low altitude go-aroundactivates the aural message “DON'T SINK, DON'T SINK” and:− Amber "GND PROX" indication on the PFD.

The audio message is only annunciated twice, unless excessivealtitude loss continues to accumulate.

Once triggered, the visual message can only be cancelled achieving apositive rate of climb relative to the original altitude. Therefore, as longas the original altitude is not crossed, any descent will trigger the auraland visual messages again. After crossing the original altitude, a newaltitude value is set every moment.

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GPWS/EGPWS MODE 3 SCHEMATIC

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MODE 4 - INSUFFICIENT TERRAIN CLEARANCE

Mode 4 provides alerts for insufficient terrain clearance with respect tophase of flight and speed. Mode 4 exists in three forms, 4A, 4B and4C.

MODE 4A

Mode 4A is active during cruise and approach with the landing gearUP.

Minimum Terrain Clearance (MTC) for “TOO LOW GEAR” messagetriggering:− Minimum: 30 ft.− Maximum: 500 ft for an airspeed equal or less than 190 KIAS.

Minimum Terrain Clearance (MTC) for “TOO LOW TERRAIN”message triggering:− Minimum: 30 ft.− Maximum: 1000 ft for an airspeed equal or higher than 250 KIAS.

If during cruise the ground is slowly getting closer and the airplane isnot in the landing configuration or during approach with anunintentional gear up landing, the aural message "TOO LOWTERRAIN" will be sounded. Once the message has been issued, anadditional 20% altitude loss is required for the issuing of a newmessage.

The amber "GND PROX" indication is displayed on the PFD.

If the airplane penetrates below the 500 ft AGL boundary with thelanding gear still up, the aural message will be "TOO LOW GEAR".Once a message is issued, an additional 20% altitude loss is requiredfor the issuing of a new message.

The visual and aural messages cease when the mode 4A is exited.

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EGPWS MODE 4A SCHEMATIC

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MODE 4B

Mode 4B is active during cruise and approach with the landing geardown and flaps in other than landing configuration.

Minimum Terrain Clearance (MTC) for "TOO LOW FLAP" messagetriggering:− Minimum: 30 ft.− Maximum: 245 ft for an airspeed equal or less than 159 KIAS.

Minimum Terrain Clearance (MTC) for “TOO LOW TERRAIN”message triggering:− Minimum: 30 ft.− Maximum: 1000 ft for an airspeed equal or higher than 250 KIAS.

If during cruise the ground is slowly getting closer and the airplane isnot in the landing configuration, or during approach with anunintentional gear up landing, the aural message "TOO LOWTERRAIN" will be sounded. Once the message is issued, an additional20% altitude loss is required for the issuing of a new message.


If the airplane penetrates below the 245 ft AGL boundary with thelanding gear down and flaps in other than landing configuration, theaural message will be "TOO LOW FLAPS". Once message is issued,an additional 20% altitude loss is required for the issuing of a newmessage.

The visual and aural messages cease when the mode 4B is exited.

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EGPWS MODE 4B SCHEMATIC

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MODE 4C

Mode 4C is active during takeoff phase or low altitude go-around witheither the landing gear or flaps in other than landing configuration,when the terrain is rising closer than the airplane is climbing.

Only in this case, the Minimum Terrain Clearance is a function of theRadio Altitude of the airplane.

Minimum Terrain Clearance (MTC) for "TOO LOW TERRAIN"message triggering:− Minimum: 30 ft.− Maximum:

− 500 ft at 667 ft or greater of radio altitude for an airspeed less or equal or less than 190 KIAS.− 1000 ft at 1333 ft or greater of radio altitude for an airspeed equal or above 250 KIAS.

If during takeoff or low altitude go-around with either the landing gearor flaps in other than landing configuration, when the terrain is risingmore steeply than the airplane is climbing, the aural message "TOOLOW TERRAIN" will be sounded.


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EGPWS MODE 4C SCHEMATIC

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MODE 5 - EXCESSIVE DEVIATION BELOW GLIDESLOPE BEAM

Mode 5 provides two levels of alerting if the airplane's flight pathdescends below the glideslope on ILS approaches.

Minimum Terrain Clearance (MTC) for "GLIDESLOPE" messagetriggering:− Minimum:

− For the Soft Alert Area, 30 ft at 2.98 dots of glideslope deviation.− For the Hard Alert Area, 30 ft at 3.68 dots of glideslope deviation.

− Maximum:− For the Soft Alert Area 1000 ft.− For the Hard Alert Area 300 ft.

The first alert occurs whenever the airplane is more than 1.3 dotsbelow the beam and is called a "soft alert" because the volume level isreduced. A second alert occurs below 300 ft radio altitude with greaterthan 2 dots deviation from glideslope and is louder or "hard".

The aural message "GLIDESLOPE" is sounded once. Follow-on alertsare only allowed when the airplane descends lower on the glideslopebeam by approximately 20%. Aural messages are soundedcontinuously once the airplane exceeds 2 dots.


The glideslope warning can be canceled by pressing the MasterCaution Button.

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EGPWS MODE 5 SCHEMATIC

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MODE 6 - CALLOUTS

Mode 6 provides aural messages for descent below predefinedaltitudes, decision height, a minimums setting and approachingminimums. Alerts for excessive roll or bank angle are also provided.

MINIMUMS CALLOUTS

The message "APPROACHING MINIMUMS" is sounded only oncewhen the airplane is 80 ft above the decision height or another targethas been reached, with the landing gear down.− Radio altitude for message triggering:

− Minimum: 90 ft.− Maximum: 1000 ft.

The message "MINIMUMS MINIMUMS" is sounded only once whenthe airplane is at decision height or another target has been reached,with the landing gear down.− Radio altitude for message triggering:

− Minimum: 10 ft.− Maximum: 1000 ft.

Visual indication of minimum target is presented on PFD.

EGPWS MODE 6 - SCHEMATICMINIMUMS CALLOUTS

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ALTITUDE CALLOUTS

The messages "FIVE HUNDRED”, “TWO HUNDRED” and“ONE HUNDRED" will be sounded when associated radio altitude hasbeen reached, with the landing gear down.

The “FIVE HUNDRED” message will only be sounded whether one ormore of the following conditions are satisfied:− ILS is not tuned or not available.− ILS is tuned in a valid signal, but with a deviation greater than 2

dots of localizer or glideslope.− If a backcourse approach is detected.

Radio altitude for message activation:− Minimum: 50 ft.− Maximum: 1000 ft.

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01 REVISION 2

BANK ANGLE CALLOUT

Minimum Terrain Clearance (MTC) for message triggering onEGPWS:− Minimum: 5 ft.− Maximum: Increases linearly from 30 ft at 10° of bank angle to 150

ft at 40° then from 150 ft at 40° up to 2450 ft at 55°, remaining constant at 55° above 2450 ft.

The aural message "BANK ANGLE, BANK ANGLE" is sounded whenthe airplane bank angle is too high or roll rate exceeds 1°/sec during allphases of flight.

The message is generated again if bank angle increases by 20%.

When roll attitude increases to 40% above the initial callout angle, thecallout will repeat continuously.

EGPWS MODE 6 - SCHEMATICBANK ANGLE CALLOUT

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EGPWS ADDITIONAL FEATURES

The EGPWS also includes the Terrain Clearance Floor, Terrain LookAhead Alerting and Terrain Awareness Display features. Airplanesequipped with EGPWS version 216 incorporates additional featureslike Peaks Mode, Runway Field Clearance Floor, Obstacle Alerting andGeometric Altitude.

TERRAIN CLEARANCE FLOOR

The Terrain Clearance Floor (TCF) provides a terrain clearancecircular envelope around the airport runway, alerting the pilot of apossible premature descent for non-precision approaches regardlessof the airplane's configuration. The TCF is active during takeoff, cruiseand final approach and is based on current airplane position, nearestrunway and radio altitude.

This alert mode complements the Mode 4 by providing an alert basedon insufficient terrain clearance even when the airplane is in thelanding configuration.

TCF alerts display “GRND PROX” on the PFD and the aural message"TOO LOW TERRAIN" sounds. This message sounds once wheninitial envelope penetration occurs and will repeat at every additional20% decrease in radio altitude. The “GRND PROX” annunciatorremains on until the TCF envelope is exited.

In the EGPWS version 216, the TCF alert provides an envelopeextension for runway sides, which is limited to a minimum value of245 ft beside the runway, within 1 NM to 2.5 NM from runway end. Thisfeature provides improved alerting when it is determined that theaircraft is landing to the side of the runway.

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TCF ALERT ENVELOPE

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TERRAIN LOOK AHEAD ALERTING

The Terrain Look Ahead Alerting provides a caution/warning level toalert the flight crew about potential terrain conflicts. The alerts arebased mainly on the airplane's current position and barometric altitudeinformation. In the event of terrain caution or warning conditions, aspecific audio alert and visual alert are triggered and the terrain displayimage is enhanced to highlight each of the types of terrain threats.

TERRAIN WARNING AND CAUTION AREAS

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When conditions are such as to generate a Terrain Caution alert(approximately 60 seconds prior to potential terrain conflict), the auralmessage "CAUTION TERRAIN, CAUTION TERRAIN" is sounded andthe amber "GND PROX" indication is displayed on the PFD. This isrepeated every seven seconds as long as the airplane is still in thecaution envelope.

When conditions have been met to generate a Terrain Warning alert(approximately 30 seconds prior to potential terrain conflict), the auralmessage "TERRAIN, TERRAIN, PULL UP" is sounded and the red"PULL UP" indication is displayed on the PFD.

The terrain image will appear automatically on the MFD when a terrainthreat event occurs.

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01 REVISION 6

TERRAIN AWARENESS DISPLAY

The EGPWS terrain display is designed to increase flight crewawareness of the surrounding terrain in varying density dots patterns ofgreen, yellow and red. These dot patterns represent specific terrainseparation with respect to the airplane. The following table relates thecolor that the terrain is displayed with its meaning:

COLOR MEANINGSolid red Warning Terrain

(Approximately 30 sec from impact).Solid yellow Caution Terrain

(Approximately 60 sec from impact).High density red dots Terrain that is more than 2000 ft

above airplane altitude.High density yellow dots Terrain that is between 1000 and

2000 ft above airplane altitude.Medium density yellowdots

Terrain that is between 500 ft (250 ftwith gear down) to 1000 ft belowairplane altitude.

Medium density green dots Terrain that is between 500 ft (250 ftwith gear down) below and 1000 ftbelow airplane altitude.

Light density green dots Terrain that is 1000 to 2000 ft belowairplane altitude.

Black Terrain below 2000 ft.

NOTE: - Terrain is not shown if its elevation is within 400 ft of runwayelevation of the nearest airport.

- To reduce clutter on the display, any terrain more than 2000 ftbelow the airplane is not displayed.

- Terrain that is not covered in the EGPWS database will bedisplayed in magenta.

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EGPWS DISPLAY COLOR CODING

EXAMPLE OF EGPWS DISPLAY ON MFD

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01 REVISION 6

PEAKS MODE

This is a feature provided only by EGPWS version 216 and, whenselected, adds additional density patterns and level thresholds to thestandard mode display levels and allows the terrain to be displayedduring the cruise phase even if it is more than 2000 ft below theaircraft.

When the Peaks Display is on, elevation numbers indicating thehighest and lowest terrain/obstacle currently being displayed areshown on the display. These elevations are expressed in hundreds offeet above sea level (MSL) with the highest elevation on top and thelowest on the bottom. In the event that there is no appreciabledifference in the terrain/obstacle elevations, only the highest value isdisplayed.

The color of the elevation value displayed matches the color of theterrain displayed.

If the aircraft is 500 ft (250 ft with landing gear down) or less above theterrain in the displayed range, the peaks color displayed will beidentical to the terrain awareness display mode, with the exception ofsea level displayed as cyan.

PEAKS PROFILE AT A LOW RELATIVE ALTITUDE

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When the aircraft is greater than 500 ft (250 ft with landing gear down)above all terrain in the displayed range, no yellow or red bands aredisplayed and low density green, medium density green and solidgreen will be displayed as a function of the highest and lowestelevations in view. Moreover, sea level elevations can be displayed ascyan to simulate water.

PEAKS PROFILE AT A HIGH RELATIVE ALTITUDE

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WARNING PRIORITIES

The GPWS/EGPWS warning priorities are listed below. Messages atthe top will start before or override a lower priority message even if it isalready in progress.

MESSAGE MODEPULL UP 1 and 2TERRAIN TERRAIN 2 and Terrain Look-AheadPULL UP Terrain Look-AheadTERRAIN 2MINIMUMS MINIMUMS 6CAUTION TERRAIN Terrain-Look AheadTOO LOW TERRAIN 4 and Terrain Clearance FloorALTITUDE CALLOUTS 6TOO LOW GEAR 4TOO LOW FLAPS 4SINKRATE 1DON'T SINK 3GLIDESLOPE 5APPROACHING MINIMUMS 6BANK ANGLE 6

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EICAS MESSAGES

EGPWS


WARNING GPWSOne GPWS envelope,associated to Modes 1 to 4,has been penetrated.

CAUTIONGPWS INOP GPWS monitor has detected an

internal failure.

TERR INOP Terrain mode is not available.

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1 - EGPWS TERRAIN SYSTEM OVERRIDE BUTTON

− When pressed, inhibits EGPWS in approach mode, thusavoiding unwanted terrain alerts in airports not covered byEGPWS database.

EGPWS TERRAIN SYSTEM OVERRIDE BUTTON

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MFD BEZEL PANEL

1 - EGPWS DISPLAY SELECTOR BUTTON− Alternate pressing will cause the MFD to toggle between the

weather radar or terrain to be displayed.− The ranges available are: 5 NM, 10 NM, 25 NM, 50 NM,

100 NM, 200 NM, 300 NM, 500 NM and 1000 NM.− When a terrain warning/caution condition exists and the terrain

is not selected on the MFD, the terrain will be automaticallydisplayed on the MFD with a range of 10 NM.

MFD BEZEL PANEL

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EGPWS DISPLAY ON MFD

1 - TERRAIN ANNUNCIATIONS

LABEL COLOR CONDITIONTERR

(Upper left corner)Cyan Lit when terrain

mode is selected.TERR FAIL Amber Lit when terrain

mode is inoperative.TERR INHIB forTerrain Inhibition

White Lit when the EGPWSterrain systemoverride button ispressed in approachmode.

TERR N/A Amber Lit when EGPWS isuncertain of theairplane's position.

TERR TEST Red Lit when the self testis activated.

2 - TERRAIN INDICATION− Displays an image of surrounding terrain in varying density dot

patterns of green, yellow and red. These dot patterns representspecific terrain separation with respect to the airplane. Thedisplay is generated from airplane altitude compared to terraindata.

3 - TERRAIN ALERT INDICATION− Indicates a terrain warning or caution condition.

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EGPWS DISPLAY ON MFD

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01 REVISION 6

DISPLAY ON PFD

EGPWS

1 - PULL UP/GROUND PROXIMITY ANNUNCIATIONS

− Label: PULL UP (red)GND PROX for Ground Proximity (amber).

− PULL UP is lit when either modes 1 or 2 have been activated intheir more critical situation.

− GND PROX is lit when ground is getting closer too fast.

EGPWS DISPLAY ON PFD

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STEEP APPROACH OPERATION

Some airplanes may be optionally equipped with Steep Approachfunctions. Steep Approaches are approach operations performed withglide slope angle above 4.4 degrees. This kind of operation implies tothe airplane a vertical speed higher than the normal, requiring meansto change the range of the EGPWS Mode 1 envelope in order to avoidnuisance messages.

The Steep Approach mode is selected by means of two pushbuttonsinstalled on the glareshield panel, one at each side. When eitherpushbutton is pressed, an internally preset mode of the EGPWSchanges the references to sound the SINK RATE and PULL UP auralwarnings.

When the airplane is in flight and the flaps are selected to 45°, theSTEEP white light illuminates on the Steep Approach pushbuttonindicating that the Steep Approach mode is available. When either theflaps are retracted to a position other than 45° or airplane lands, theSTEEP white light extinguishes indicating that the Steep Approachmode is no longer available.

The pushbutton lower portion has two status lights, amber and green.The green light indicates that the Steep Approach mode is engagedand the amber light indicates a failed condition.

If the amber light turns on, it indicates that the Steep Approach mode isfailed and steep approach operations must not be performed. In thissituation, the Steep Approach mode may or not be engaged and theairplane must land in an airport that not requires steep approachoperation. The pilot must monitor the vertical speed during theapproach and landing phases.

In flight, with the STEEP inscription illuminated if the Steep Approachpushbutton is pressed, the green light illuminates to indicate that theSteep Approach mode is engaged. If the green light does notilluminate, the Steep Approach mode is not engaged and the steepapproach operation must not be performed.

The Steep Approach mode is deselected pressing the pushbutton orthrough automatic deselection. An automatic deselection of the SteepApproach mode is performed when:− Airplane on the ground;− Flaps setting other than 45°.

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01 REVISION 7

STEEP APPROACH MODE PUSHBUTTON

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STEEP APPROACH BUTTON

LIGHT INDICATION MODE DESCRIPTIONSTEEP Illuminates in white color when the airplane is in

the air and the flaps are in 45°. This means thatthe Steep Approach mode is available.

GREEN LIGHT Illuminates when the button is pressed with theSTEEP light illuminated. This means that theSteep Approach mode is engaged.With the STEEP light illuminated, if the greenlight does not illuminates when the pushbuttonis pressed, means that the Steep Approach isnot engaged; in this case, do not perform SteepApproach operations.

AMBER LIGHT The Steep Approach mode is failed. Do notperform Steep Approach operations. In thissituation, the Steep Approach mode may or notbe engaged and the airplane must land in anairport that not requires steep approachoperation. The pilot must monitor the verticalspeed during the approach and landingphases.

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WINDSHEAR DETECTION AND ESCAPEGUIDANCE SYSTEMThe EMB-135BJ is equipped with an additional warning systemdedicated to windshear detection. The system provides visual andaural alarms to warn pilots of a windshear occurrence, as well as themost appropriate maneuver to recover from such phenomenon.The Windshear Detection function is performed by the EGPWScomputer, which also performs ground proximity warning functions.The Windshear Escape Guidance is a Flight Director mode providedby the avionics package.

WINDSHEAR GENERAL INFORMATION

Windshear is a sudden change in wind direction or speed, normallycaused by thunderstorms, frontal systems or any topographical featurethat may affect the wind flow (e.g. hills, mountains, lakes, seas,...).

Due to ground proximity, the most hazardous phases of flight regardingwindshear encounters are takeoff, approach and landing. On awindshear, wind may shift from a tailwind to a headwind or to adowndraft or updraft. The consequences may be an abrupt change inairspeed, lift and altitude, upwards or downwards, according to shiftingdirection. Although quick, windshear is not instantaneous, which maylead pilots to correction attempts in the wrong manner. For instance,an airplane facing a headwind after takeoff, appears to have goodperformance, characterized by high airspeed, which drives the pilot intorotating the airplane to a pitch higher than usual. When thethunderstorm core is reached, wind shifts to a downdraft and airspeeddecreases, as well as vertical speed. The pilot’s natural reaction is tolower the airplane’s nose in an attempt to maintain airspeed. Furtherahead, wind shifts to tailwind component, resulting in a dramaticairspeed reduction with the nose already down. Under such scenario, itis very difficult to maintain a positive rate of climb.If the takeoff or landing can not be delayed, the correct action is toincrease airspeed before being subjected to windshear encounter andto consider flying near stall speeds with high angle of attack ifnecessary to regain altitude.

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WINDSHEAR DETECTIONThe windshear detection system is designed to identify the presence ofsevere windshear phenomenon and to provide timely warnings andadequate flight guidance for approach, missed approach, takeoff andclimb out.

The windshear computer exchanges data with AHRS, ADC, SPS,Radio Altimeter and IC-600s. The system continuously searches forany windshear clue, and then signals the PFD and aural warning unit toprovide the appropriate indications.

Windshear Caution alerts are given if the windshear consists of anincreasing headwind (or decreasing tailwind) and/or severe updraft,which may precede an encounter with a microburst. Windshearcautions activate the Windshear Caution (WDSHEAR) amberindications on the upper left corner of both PFDs and an auralmessage “CAUTION WINDSHEAR” is also triggered. WindshearCaution indications remain on for as long as the airplane remainsexposed to an increasing headwind and/or updraft condition in excessof the alert threshold.Windshear Warnings are given if the windshear consists of adecreasing headwind (or increasing tailwind) and/or severe downdraft.Windshear warnings activate the Windshear Warning (WDSHEAR) redindication on both PFDs and trigger an aural message “WINDSHEAR,WINDSHEAR, WINDSHEAR”. This message will not be repeatedunless another, separate, severe windshear event is encountered.Windshear Warning indications remain on for as long as the airplaneremains exposed to a decreasing headwind and/or downdraft inexcess of the alert threshold. The threshold is adjusted in function ofavailable climb performance, flight path angle, airspeeds significantlydifferent from normal approach speeds and unusual fluctuations inStatic Air Temperature (typically associated with the leading edges ofmicrobursts).

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WINDSHEAR ESCAPE GUIDANCE MODE

The Windshear Escape Guidance Mode is used to minimize altitudeand speed loss during a windshear encounter. The strategy is to keepthe airplane airborne until the windshear conditions subside or areexited.The Windshear Escape Guidance Mode provides pitch command torecover from a windshear encounter. The amplitude of the pitchcommand will depend upon the airplane’s performance and windshearseverity and phase.

The Windshear Escape Guidance is a Flight Director mode engagedunder the following conditions:

− Manually, by pressing the Go Around Button while a windshearcondition (increasing/decreasing performance) is detected;

− Automatically, when in Go Around or Takeoff Mode and a windshearcondition (increasing/decreasing performance) is detected;

− Automatically, when Thrust Levers Angle is above 78° and adecreasing performance windshear is detected (windshear warning).

When the windshear escape guidance mode is engaged a green“WSHR” indication is displayed on both PFDs in the Vertical Mode fieldand a “ROLL” indication is displayed in the Lateral Mode field.Whenever the Windshear Escape Guidance mode is engaged, thePitch Limit Indicator (PLI) symbol is drawn directly on the AttitudeDisplay Indicator portion of the PFD. The PLI represents the remainingangle of attack margin before Stick Shaker triggering.

All other Flight Director modes are canceled and the following verticalmodes are inhibited when a caution or warning windshear condition ispresented:

− Altitude Preselect Mode, Go Around and Takeoff.

No lateral modes are inhibited while in the vertical mode of WSHR.

The Windshear Escape Guidance mode is designed to meet thefollowing requirements, in the listed order of priorities:

− Prevent the airplane from stalling;

− Prevent the airplane from descending;

− Prevent the airplane from exceeding VMO.

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The Windshear Escape Guidance Mode incorporates three controlsub-modes:

− Alpha Sub-mode - The airplane can be commanded to descend inorder to maintain airspeed when approaching stall conditions. If theflight path angle control results in an angle of attack beyond the stickshaker triggering angle, the windshear control law can keep theairplane angle of attack below the stick shaker threshold.

− Gamma Sub-mode - The airplane can be prevented fromdescending by commanding a positive flight path angle. A nominalflight path angle is used to allow an airspeed raise during anincreasing performance windshear, in anticipation of a decreasingperformance windshear, and also to minimize altitude loss during adecreasing performance windshear.

− Speed Target Sub-mode - The airplane is allowed to climb in order toexchange excessive kinetic energy for potential energy. If the controlof the flight path angle results in an excessive speed increase, thewindshear control law maintains the airplane indicated airspeed atthe target speed.

The Windshear Escape Guidance mode will be canceled if any of thefollowing conditions occur:

− FLC, VS, SPD or ALT Mode is selected;− Invalid IRS data;− Invalid ADC data;− Invalid Stall Protection Computer (SPC);− Radio Altitude greater than 1500 ft.

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EICAS MESSAGE


CAUTION WINDSHEAR INOPWindshear detection andescape guidance system isinoperative.



1 - WINDSHEAR INDICATION− Indicates that a windshear has been detected.− Color: amber or red depending on windshear severity.

2 - ESCAPE GUIDANCE MODE ENGAGEMENT ANNUNCIATION− Indicates the Windshear Flight Guidance Escape Mode

engagement.

3 - PITCH LIMIT INDICATOR− Refer to Stall Protection System indicators in section 2-04-25.

4 - FLIGHT GUIDANCE INDICATION− Indicates the appropriate pitch to be attained, during a

windshear occurrence.

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(V-BAR AND CROSS-BAR FORMAT)

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TRAFFIC AND COLLISION AVOIDANCE SYSTEM

GENERAL

The EMB-135BJ is equipped with a Traffic and Collision AvoidanceSystem (TCAS), which provides the flight crew with an indication ofpossible in-flight traffic conflict. The system is based upon transpondersignals and provides visual and aural warnings, as well asrecommended evasive action.

The EMB-135BJ may be equipped with TCAS software version 6.04A(TCAS II and TCAS 2000) or with TCAS software version 7.0(TCAS 7).

The TCAS 2000 presents the same operational characteristics of theTCAS II.

The TCAS 7 presents the following differences when compared to theTCAS II or TCAS 2000:− The altitude separation thresholds for issuing Traffic Advisory (TA)

and Resolution Advisory (RA) between FL300 and FL420 arereduced for compatibility with RVSM flight operations.

− The thresholds for issuing RA for airplanes closing in altitude arereduced between the FL200 and FL420.

− Reduction in the numbers of RA eliminating those airplanes that areexpected to pass with sufficient horizontal range separation.

− Allows RA direction reversion, i.e, change a CLIMB to a DESCENTand vice-versa in coordination with another TCAS equippedairplane.

− Introduction of three additional RA.− Different set points and range of actuation, as presented in the text

below.

SYSTEM DESCRIPTION

The TCAS was developed to provide crew awareness regardingpossible conflicting air traffic situations. Besides providing awareness,TCAS also displays to the flight crew the recommended verticalmaneuver to avoid conflicting traffic. TCAS does not providerecommendations for horizontal maneuvers.

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CAUTION: PRIMARY RESPONSIBILITY FOR EVASIVE ACTIONLIES WITH THE FLIGHT CREW AND ANY ACTIONMUST ALWAYS BE PRECEDED BY A VERY CAREFULEVALUATION OF THE SITUATION.

The TCAS computer receives data from the installed transponders,radio altimeters and air-ground sensor. The signals transmitted bysurrounding airplanes inform their altitude, bearing and identification,thus making it possible to track any traffic that could enter theairplane’s protection zone. Based on such data, the TCAS calculatesthe predicted path of each intruder airplane, determining whether ornot it may become a target. To determine that, an alert zone isestablished, based on separation and speeds of both airplanes. Thesize of the alert zone is not distance-based but, rather, is based ontime. Therefore, the caution area corresponds to the volume in spacewhere a conflict is expected to occur in 35 to 45 seconds, if no action istaken. A warning area corresponds to an imminent conflict in thefollowing 20 to 30 seconds. Such time is calculated by dividingdistance between airplanes by their closure rate.

To inhibit the issuing of undesired warnings that constitute a nuisanceeffect, the system incorporates a series of protections. These applyduring approaches to crowded airports, to increase protection againstslow closure rates, and to prevent airplanes below 180 ft (380 ft forTCAS 7), which are about to land or have just taken-off, from creatinga nuisance.

When an airplane is tracked by the TCAS, the system periodicallyinterrogates the intruder’s transponder. The exchange of data betweentwo subsequent transmissions makes it possible to obtain the distanceto the intruder and its altitude, and to predict its path.

If the predicted path of the intruder enters the airplane’s alert area, twokinds of alerts may be generated. If the area to be penetrated is thecaution area, a Traffic Advisory (TA) is generated. Pilots are thenrequested to visually locate the intruder and perform the requiredpreventive action.

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If the warning area is penetrated a Resolution Advisory (RA) isgenerated, as well as the corrective action that must be taken to permitthe greatest possible separation at the Closest Point of Approach(CPA). Sometimes, the recommended action may lead to crossing ofthe intruder’s flight level or may change during the maneuver. Thissituation may occur when the calculation indicates that this is the bestway to achieve the greatest possible separation at the CPA. For bothadvisory cases, a symbol is presented in the MFD to indicate theintruder’s relative position, altitude and danger level. A voice messageis generated to help the pilots in taking the most suitable action. ThePFD provides indication of the recommended vertical speed to clearthe conflict. A voice message may be generated to warn the pilot intomonitoring the VSI on the PFD. When TCAS computations indicatethat the traffic has been cleared, a voice message advises pilots thatthere is no longer a conflicting situation. In this condition, if no other TAor RA is on course, the intruder’s indication changes, indicating that itis a safe nearby traffic.If the intruder is also equipped with a TCAS, maneuvers arecoordinated between both airplanes. If the intruder is only equippedwith a transponder, the system may still indicate its position, providedits transponder is at least mode C. For airplanes equipped with modeA transponder, only Traffic Advisories may be generated.

CAUTION: THE TCAS CAN ONLY GENERATE RESOLUTIONADVISORIES FOR INTRUDERS EQUIPPED WITHRESPONDING MODE S OR MODE C TRANSPONDERS.TRAFFIC ADVISORIES CAN BE GENERATED FORAIRPLANE WITH OPERATIVE MODE S, MODE C ORMODE A TRANSPONDERS. THE TCAS PROVIDES NOINDICATION OF AIRPLANE WITHOUT OPERATINGTRANSPONDERS.

System options may be monitored and set through the RMU. Adedicated window is provided, presenting which TCAS display is beingcontrolled, its range and altitude band. A RMU page permits togglingbetween options. Controls allow selection of different ranges, eitherhorizontal and vertically, as well as changing the way some parametersare presented.

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(*)

(*) 380 ft for TCAS 7.

TCAS PROTECTED AREAS

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TCAS VOICE MESSAGES

TYPE MESSAGE MEANING REMARKS

TA

TRAFFIC,TRAFFIC

An intruder is expected toenter the collision area in35 to 45 seconds. Anindication of it has justbeen displayed on theMFD.

− For TCAS II, seeNOTE 1.

− For TCAS 7, all TAare inhibited below500 ft AGL.

MONITORVERTICAL SPEED

Vertical speed is changingto a non-recommendedvalue.

PREVENTIVE

ADJUST VERTICALSPEED, ADJUST

Vertical speed has to beadjusted to therecommended valueindicated on the VSI.

TCAS 7 only.

RA MAINTAINVERTICAL SPEED,MAINTAIN

Maintain the vertical speedindicated on the VSI.

TCAS 7 only.

MAINTAINVERTICAL SPEED,CROSSINGMAINTAIN

Maintain the vertical speedindicated on the VSI.During climb or descent,airplane will crossintruder’s flight level.

TCAS 7 only.

CLIMB Climb at the vertical speedindicated on the VSI toclear the possible conflict.

CORRECTIVE

DESCEND Descend at the verticalspeed indicated on the VSIto clear the possibleconflict. Vertical Speed willbe 1500 ft/min or greater.

See NOTE 1.

RA REDUCE CLIMB Reduce climb speed toclear the possible conflict.

Not valid for TCAS 7.

REDUCEDESCENT

Reduce descent speed toclear the possible conflict.

− See NOTE 1− Not valid for TCAS 7.

CLIMB,CROSSINGCLIMB

Climb at the indicatedvertical speed on the VSIto clear possible conflict.During climb, airplane willcross intruder’s flight level.

(Continued)

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TYPE MESSAGE MEANING REMARKSDESCEND,CROSSINGDESCEND

Descend at the indicatedvertical speed on the VSIto clear possible conflict.During descend, airplanewill cross intruder’s flightlevel.

See NOTE 1.

INCREASECLIMB

Climb speed has to beincreased to therecommended value toclear the possible conflict.

Vertical Speed must be2500 ft/min or greater.

CORRECTIVERA

INCREASEDESCENT

Descent speed has to beincreased to therecommended value toclear the possible conflict.Vertical Speed must be2500 ft/min or greater.

− For TCAS II, thismessage is inhibitedbelow 1450 ft AGL.

− For TCAS 7, thismessage is inhibitedbelow 1450 ft AGLwhile descending andbelow 1650 ft AGLwhile climbing.

CLIMB, CLIMBNOW!

After a descent advisory,TCAS detected a changingsituation that requires theneed to climb.

DESCEND,DESCEND NOW!

After a climb advisory,TCAS detected a changingin situation that requiresthe need to descend.

See NOTE 1.

CLEAR OFCONFLICT

The possible conflict hasbeen cleared. Message ispresented only if intruder’stransponder signal is valid.

Not presented if theintruder track or altitudeinformation is lost.

NOTE: 1) Inhibited below 1000 ft AGL while descending and below1200 ft AGL while climbing.

2) All RAs are inhibited below 400 ft AGL while descendingand below 600 ft AGL while climbing.

3) For TCAS II, RA messages are repeated three times (one-word messages) and twice (two-word messages). ForTCAS 7, all RAs are repeated twice.

4) TA message sounds once.

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RMU RADIO PAGE - ATC/TCAS WINDOW

Refer to Section 2-18 - Navigation and Communication for furtherdetails on RMU controls.Refer to RMU ATC/TCAS Control Page in this Section for furtherdetails on TCAS controls.

1 - TRANSPONDER OPERATING MODE− Allows selection of TCAS modes:

− TA ONLY - TCAS traffic advisory mode is selected.− TA/RA - TCAS traffic advisory and resolution advisory modes

are selected.− Refer to Section 2-18 - Navigation and Communication for

further details.

2 - TCAS CONTROL SIDE IDENTIFICATION− Indicates which TCAS display (MFD 1 or 2) is being controlled

through that RMU. The selection of TCAS DSPY 1 or 2 isaccomplished through the cross-side transfer button when theyellow cursor box is placed on this field.

− Color: white for on-side TCAS display and magenta for cross-side.

3 - TCAS RANGE DISPLAY − Displays the selected TCAS range value.− Color: green− Possible selections are 6, 12, 20, 40 NM. Airplanes equipped

with TCAS 7 also allow 80 and 100 NM selection.

4 - TCAS ALTITUDE BAND INDICATION− Indicates the TCAS altitude band according to selected TCAS

mode.− NORMAL (green) - With the TA display set to AUTO the

operational TCAS altitude band will be from 1200 ft below to1200 ft above the airplane. With the TA display set toMANUAL the operational TCAS altitude band will be from2700 ft below to 2700 ft above the airplane.

− ABOVE - The operational TCAS altitude band will be –2700 ft to +7000 ft.

− BELOW - The operational TCAS altitude band will be –7000 ft to +2700 ft.

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1 - INTRUDER ALTITUDEREL (green) - Intruder’s altitude is displayed as a relative altitude to

the airplane. Value is preceded by a plus or a minussignal, depending on whether the intruder is above orbelow the airplane.

FL (cyan) - Intruder’s altitude is displayed as its flight level. Thisselection automatically reverts to REL after 20seconds.

2 - TA DISPLAYAUTO - Traffic is displayed only when a TA or RA condition

exists.MANUAL - All traffic detected by the system is displayed.

3 - FLIGHT LEVEL 1/2− Displays the transponder-encoded altitude and the air data

source.Refer to transponder description (Section 2-18 – Navigation andCommunication).

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TCAS TEST

The TCAS self-test is activated through the RMU TST button and maybe performed on the ground or in flight. TCAS will operate normally ifnot tested.

To test the system proceed as follows:− On the RMU radio page, set the ATC/TCAS window to the TA/RA

mode. On the MFD, set TCAS mode.− Press and hold for 7 seconds the RMU TST button.− A white TCAS TEST message will be presented on the MFDs and

PFDs.− A TCAS TEST aural warning will sound.− The Master Warning lights will flash.

− The MFDs show a traffic test parttern, which permits the checkingof each of the existing intruder symbols, i.e., a hollow bluediamond, a solid blue diamond, a solid amber circle and a solidred square.

− On the PFDs, the VSI shows red and green arc zones.− At the end of the test, the RMU shows a green TCAS PASS

message and a TCAS TEST PASS aural warning will sound.

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MULTI FUNCTION DISPLAY

1 - INNER RANGE RING− Displayed around airplane symbol to indicate a 2 NM range.− Removed if outer range indicates distance above 20 NM.

2 - OUTER RANGE RING− May be selected up to 40 NM. Airplanes equipped with TCAS 7

allow selection up to 100 NM.

3 - NO BEARING ADVISORIES INDICATION− Indicates data related to a detected intruder, whose bearing

cannot be determined.− Up to two lines may be displayed indicating the kind of advisory,

its distance, relative altitude and whether it is climbing ordescending in excess of 500 ft/min.

− Colors: No bearings RAs: red.No bearings TAs: amber.

4 - PROXIMATE TRAFFIC INDICATION− Indicated by a solid cyan diamond.− Represents any airplane within 6.5 NM horizontally and 1200 ft

vertically, but whose path is not predicted to penetrate theCollision Area.

5 - INTRUDER’S VERTICAL MOVEMENT− Indicated by an arrow next to the symbol that indicates if the

intruder is climbing or descending in excess of 500 ft.− Color: Same as of the associated symbol.

6 - INTRUDER’S ALTITUDE− Indicated by a solid two-digit number below or above the

intruder’s symbol.− Color: Same as of the associated symbol.− Normal presentation is relative altitude, which displays the

intruder’s relative altitude in hundreds of feet. A plus or minussignal indicates if the intruder is above (+) or below (–) theairplane.

− Two question marks (“??”) are displayed if the intruder’s relativealtitude is greater than 9900 ft, below or above.

− If intruder is below the airplane, intruder’s altitude is displayedbelow its symbol.

− If intruder is above the airplane, intruder’s altitude is displayedabove its symbol.

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7 - RESOLUTION ADVISORY INDICATION− Indicated by a solid red square.

8 - TRAFFIC ADVISORY INDICATION− Indicated by a solid amber circle.

9 - OTHER TRANSPONDER REPLYING TRAFFIC INDICATION− Indicated by a hollow cyan diamond.− Indicates other airplanes equipped with transponder within the

specified range and 2700 ft of vertical separation.− Not displayed if a TA or RA is in process.

10 - OUT OF RANGE INTRUDER− Indicates detected intruders that are out of display range.− Indicated as half the associated symbol.

11 - INTRUDER’S ALTITUDE MODE INDICATION− Indicates whether the selected intruder’s altitude is relative or

flight level.

12 - TCAS BAND SELECTED− Indicates whether the selected band for TCAS is below or

above.

13 - TCAS MODE ANNUNCIATIONS− Indicates current TCAS mode.− Colors and labels are as follows, in the order of priority:

− TCAS TEST - white− TCAS OFF - white− TCAS FAIL - amber− TA ONLY - white− TCAS - white− TCAS AUTO - white

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For further information on Vertical Speed Indicator, refer to Section2-17 – Flight Instruments.

VSI− Indicates the recommended vertical speed to avoid a possible

conflict.− Green range - displayed along the scale, indicates the range of

vertical speeds to be attained to avoid a conflict situation.− Red range - displayed along the scale, indicates the range of

vertical speeds prohibited for the current situation.− Green range may be displayed together with the red range or

split in two parts, depending on situation.− Red range may be displayed alone, together with the green

range, or split in two parts, depending on the situation.


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SECTION 2-05

ELECTRICAL

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-05-05 ..01DC System ......................................................................... 2-05-05 ..02

DC System Protection .................................................... 2-05-05 ..04External Power Source................................................... 2-05-05 ..05Batteries ......................................................................... 2-05-05 ..06Backup Battery ............................................................... 2-05-05 ..07Generators...................................................................... 2-05-05 ..07APU Starter-Generator ................................................... 2-05-05 ..08Electrical Distribution Logic ............................................ 2-05-05 ..09Ground Service Bus ....................................................... 2-05-05 ..10Avionics Master .............................................................. 2-05-05 ..11AC System (if applicable) ............................................... 2-05-05 ..12

EDL Configurations and Diagrams .................................... 2-05-10 ..01Abnormal Operation Configurations ............................... 2-05-10 ..01Normal, Abnormal and Emergency

Operation Diagrams................................................. 2-05-10 ..13EICAS Messages ............................................................... 2-05-15 ..01Controls and Indicators ...................................................... 2-05-15 ..03

Electrical System Panel.................................................. 2-05-15 ..03MFD Electrical Page....................................................... 2-05-15 ..10

Circuit Breaker Panel andLoad Distribution ...................................................... 2-05-20 ..01

Circuit Breaker Panel...................................................... 2-05-20 ..01DC Bus Load Distribution ............................................... 2-05-20 ..10

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GENERALThe electrical power system supplies AC and DC voltage to all loadsduring normal or emergency operation.Two different types of sources provide electrical power supply:

− DC Power− AC Power

NOTE: Airplanes S/N 863 and on are not supplied with AC power.

The DC power system supplies 28 V DC for all airplane electrical loadsand recharges the batteries. It is the primary electrical power supplysystem. The DC power system is comprised of:− Four independent generators (28 V DC/400 A/engine driven).− One APU starter-generator (28 V DC/400 A).− Two Nickel-Cadmium batteries (24 V DC/44 Ah/1 hour rate).− One lead-acid backup battery (24 V DC/5 Ah/10 hour rate).− External power source.

AC power is supplied by one 250 VA/400 Hz single-phase staticinverter, which converts 28 V DC into 115 V AC.

A dedicated page on the MFD (electrical page) provides, on request,information regarding system configuration, load and voltageconditions as well as battery temperatures. Furthermore, warning andcaution messages are presented on the EICAS to indicate an electricalsystem failure.


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DC SYSTEMThe 28 V DC electrical power system automatically controls powercontactors, fault protection, load shedding and emergency systemoperation. This reduces pilot workload during normal operation,external power supply or system failures. The Electrical DistributionLogic (EDL) and Generator Control Units (GCU) perform systemmanagement. Detected system failures are automatically isolated,causing some bus(es) to be deenergized.

Under normal operation, the electrical DC system is divided intoisolated left and right electrical networks. The left network includesgenerators 1 and 3, driven by engine 1. Operated in parallel,generators 1 and 3 are connected to DC BUS 1 to supply ESSENTIALDC BUS 1, SHED DC BUS 1 and HOT BUS 1. Battery 1 is charged bythe generators connected to DC BUS 1. Similarly, generators 2 and 4power the right network and are driven by engine 2.

Both networks are interconnected through Bus Tie Contactors (BTC) incase of operation with less than four generators. APU generator mayreplace any inoperative generator, or may be used before enginestarting when the APU generator or Ground Power Unit (GPU) maysupply the electrical system.

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DC SYSTEM PROTECTION

The system monitors generators current and voltage to the electricallysupplied equipment to protect it from a control unit failure, such as anovervoltage or a bus failure. If an overvoltage is detected, theassociated GCU deenergizes the generator, disconnecting it from thebus.

A bus failure produces an overcurrent condition to one or moregenerators. Upon sensing this overcurrent, the GCU isolates thesystem networks, opening the BTCs. If any generator remainsoverloaded due to the failure, it is then deenergized and disconnectedfrom the bus.

As long as the generator current exceeds 400 A, a caution message ispresented on the EICAS, indicating that manual load shedding isrequired. If no action is taken, the system will be isolated and somebuses may be deenergized.

System protections are designed so that normal transients will notcause generators to be disconnected from the bus inadvertently.

Resetting of the generator after a failure is accomplished by releasingthe associated Generator Button and then pressing it again.

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EXTERNAL POWER SOURCE

The Ground Power Unit (GPU) is connected to the airplane through anexternal receptacle. The GPU supplies 28 V DC to the load buses forground operation and APU starting, independently of the internalbatteries.

The GPU has priority over any battery and generator when energizingthe airplane. Thus, the generators and the batteries cannot operate inparallel with the GPU.

The GPU Button, located on the overhead panel, controls the Externalpower supply. As soon as the Ground Power Unit is energized,properly connected to the airplane receptacle, ready to supply powerbut not connected to the buses, the GPU AVAIL inscription illuminateson the GPU Button. A identical inscription above the GPU receptaclesimultaneously illuminates.

When GPU AVAIL is illuminated and the batteries are not connected tothe buses, only the GROUND SERVICE BUS is supplied through theexternal power supply. When the GPU Button is pressed, the GroundPower Contactor (GPC) will close, allowing the external power to feedthe load buses. When the external power comes on line, the GPUAVAIL inscription on the GPU Button extinguishes itself and the whitestripe on the button lower half illuminates.

An overvoltage circuit isolates the GPU from the airplane’s electricalbuses if the GPU voltage is incorrect. External power inverse polarityprotection is also provided. To reset the system, release the GPUbutton and then press it again. If the GPU overvoltage persists, GPCwill be kept open.

The external power voltage can be monitored through the electricalpage, on the MFD. The electrical system page shows the GPU boxand its voltage. The GPU voltage indication is removed in flight.


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BATTERIES

Two 24 V DC, 44 Ampere-hour, nickel-cadmium batteries supplyessential loads in case of an in-flight failure of all generators or if bothengines are shut down simultaneously and the APU is not available.Both batteries can supply at least 40 minutes of power for essentialloads in an all-generator-failure condition.

During normal operation, Battery 1 is connected in parallel withgenerators 1 and 3 (network 1). Battery 2 is connected in parallel withgenerators 2 and 4 (network 2). Battery 2 also supplies power for APUstarting.

During APU starting, battery 1 is isolated from the load buses. Whilebattery 2 provides power for APU start, battery 1 provides stableelectrical power to the equipment that can be adversely affected byvoltage transients.

A selector switch on the overhead panel controls each battery. Whenset to the AUTO position, battery contactors (BC 1 or BC 2) actuationis controlled according to the Electrical Distribution Logic (EDL). Whenthe GPU is connected, the battery contactors open so that only theGPU can supply the load buses. When on the ground, with thebatteries as the only electrical power source, EDL deenergizes theshed buses for battery conservation. When the battery selector knob isswitched to the OFF position, the battery contactor opens, isolating thebattery from the system.

The batteries are installed in the battery compartment, on the left sideof the airplane nose section. They are ventilated in flight by forcedairflow to prevent overheating. Temperature sensors installed in eachbattery provide temperature indication to the MFD. If battery internaltemperature rises above 70°C, a warning message is presented on theEICAS. If a battery is isolated from the load buses, a caution messageis displayed on the EICAS.

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BACKUP BATTERY

A 24 V DC, 5 Ampere-hour sealed lead-acid battery provides stabilizedpower for operation of the GCUs protective function, even in case ofshort-circuit, when system voltage may drop near zero volts.

The Backup Battery Button, on the overhead panel controls the backupbattery. Pressing the button when the Battery 1 or 2 Selector Knob isset to the AUTO position connects the backup battery to the electricalsystem for charging. If the Backup Battery Button is released, a cautionmessage is displayed on the EICAS.

GENERATORS

The primary source of electrical power are the four 28 V DC, 400Amperes, independent engine-driven brushless generators, twoinstalled on each engine accessory gearbox.

Each generator is automatically controlled and protected by adedicated Generator Control Unit (GCU), provided the GeneratorControl Button on the overhead panel is pressed.

The generators will come on line when engine speed stabilizes above56.4% N2. If a failure occurs and the Generator Line Contactor (GLC)opens, a reset may be attempted once by releasing the associatedGenerator Control Button and then pressing it again.

Anytime the Generator Line Contactor is inadvertently opened orgenerator current is above 400 A, a caution message is displayed onthe EICAS. The generator voltage and current can be monitoredthrough the electrical page, on the MFD.


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APU STARTER-GENERATOR

A 28 V DC, 400 Amperes, APU-driven starter-generator supplieselectrical power during ground operation or in flight, as an alternativesource of electrical power. The APU starter generator is controlled andprotected by its dedicated GCU.

The APU Generator Button, on the Electrical System Panel, must bepressed for normal operation. The APU line contactor is actuated onand off by APU speed. If a failure occurs on the APU generator, a resetmay be attempted releasing the APU Generator Button and pressing itagain. Only one reset may be attempted.

The APU generator, when operating, is connected in parallel with thegenerators supplying DC Bus 2. If needed, the APU generator canreplace an inoperative left network generator. After starting, and withan engine driven generator inoperative, the APU generatorautomatically replaces the inoperative generator.

Three electrical sources may be used to power an APU start: groundpower unit, battery 2 or battery 2 assisted by the main generators.Battery 1 cannot be used for APU starting. Instead, it is isolated fromthe load buses to provide stable electrical power to supply equipmentthat may be affected by voltage fluctuation.

During starting, the APU Starting Contactor (ASC) is closed, allowingthe APU starter-generator to operate as a starter, energized throughthe Central DC Bus. When the APU starting cycle is completed, theASC opens. A caution message is displayed on the EICAS if the ASCdoes not open.

At 95% RPM plus seven seconds, the APU starter generator isavailable to supply electrical power to the system. In this condition, theAPU Line Contactor (ALC) is closed, connecting the APU startergenerator to the load buses. If the ALC does not close due to contactorfailure or button not pressed, a caution message is displayed on theEICAS.

The APU starter generator voltage and current may be monitored onthe MFD.

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ELECTRICAL DISTRIBUTION LOGICMany different configurations are available in the Electrical DistributionLogic (EDL) to suit any particular situation. The EDL’s architecture issymmetrical and the operational logic sequence for EDL 1 is the sameas for EDL 2. EDL 1 is composed of DC Bus 1, Shed DC Bus 1,Essential DC Bus 1 and Hot Bus 1. The EDL 2 is composed of DC Bus2, Shed DC Bus 2, Essential DC Bus 2 and Hot Bus 2.

The Central DC Bus primary function is to connect the APU generatoror GPU to the load buses through the Bus Tie Contactors (BTC). TheCentral DC Bus also provides bus interconnections in case ofsymmetrical configuration, such as generators failure or engineshutdown.

EDL logic differs depending on whether the airplane is on the groundor in flight. In flight, some buses are deenergized, depending on thepower source available.

On the ground, all the DC buses are energized if at least one of thefollowing conditions occurs:− At least three generators are on.− The GPU is on and connected to the airplane.− At least one generator is on, and the Shed Buses Selector Knob is

set to OVRD position.

The DC distribution table below shows the Electrical Distribution Logicconfiguration according to the conditions of the generators.

DC DISTRIBUTION TABLE

CONDITION RESULTS

4 or 5 Generators On Two isolated left and right electricalnetworks with all buses energized.

3 Generators OnBoth electrical networks interconnectedthrough Bus Tie Contactors with all busesenergized.

1 or 2 Generators OnBoth electrical networks interconnectedthrough Bus Tie Contactors with shedbuses deenergized.

Loss of all Generators Batteries to supply the Essential Buses(in-flight condition only).


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01 REVISION 1

GROUND SERVICE BUSThe Ground Service Bus is energized by connecting the GPUconnector to the airplane receptacle and provided the batteries andgenerators are not connected to the buses (GPC, BC 1 and BC 2 areopen).

The Ground Service Bus supplies electrical power for airplaneservicing and maintenance while on the ground. It functionsindependently of the Electrical Distribution Logic and does not energizeall electrical distribution buses.

The following lights will be powered by the Ground Service Bus:− Passenger cabin lights;− Lavatory lights;− Galley lights;− Courtesy/stairs lights;− Cockpit dome lights;− Baggage/service compartment lights.

GROUND SERVICE SCHEMATIC

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AVIONICS MASTER

The avionics master system allows manual disconnection of somenavigation and communication equipment from the load buses. Thisprevents undesirable voltage transients during APU starting on theground.

The avionics master system consists of six buses: Avionics SwitchedDC Buses 1A, 1B, 2A, 2B and Avionics Switched Essential DC Buses1 and 2. These buses are supplied by their associated DC buses. TwoAvionics Master Buttons, located on the overhead panel, controlswitching the buses.

AVIONICS MASTER SCHEMATIC


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01 REVISION 7

AC SYSTEM (APPLICABLE TO AIRPLANES UP TOS/N 854)

One 250 VA/400 Hz single phase static inverter converts 28 V DCelectrical power into 115 V AC for airplane systems requiring ACpower. The avionics system is the primary user of AC power.

The inverter is power supplied by the DC Bus 1 and controlled by theAC Power Button, on the overhead panel. If DC Bus 1 is energized andthe AC Power Button is pressed, the 115 V AC BUS is automaticallyenergized. If the DC Bus 1 is deenergized, the inverter becomesinoperative.

To reduce pilot workload, the AC Power Button should remain pressed,even after engine shutdown. If the AC Power Button is released, astriped bar illuminates to indicate that the button is out of normaloperating condition.

During normal airplane operation, if 115 V AC BUS is deenergized, acaution message is displayed on the EICAS. An inverter reset may beattempted through the AC Power Button, by releasing and thenpressing it again.

Under electrical emergency conditions the inverter stops the operation.

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MANUAL

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AC GENERATION AND DISTRIBUTION SCHEMATICS FORAIRPLANES UP TO S/N 854


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01 REVISION 1


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JUNE 20, 2002 2-05-10 Page

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01

ELECTRICAL DISTRIBUTION LOGIC (EDL)CONFIGURATIONS AND DIAGRAMS

ABNORMAL OPERATION CONFIGURATIONSFor the Electrical Distribution Logic configurations presented here, theinitial control positions on the Electrical System Panel are the following:− Generator Buttons pressed;− GPU Button released;− Battery Selector Knobs set to AUTO position;− Essential Power Button released;− Bus Tie Selector Knob set to AUTO position;− Shed Buses Selector Knob set to AUTO position;− Backup Battery Button pressed;− Avionics Master Buttons pressed.

NOTE: - All abnormal conditions considered below are in-flightconditions.

- In the schematic configurations, the continuous boxesindicate energized buses while dashed boxes indicatedeenergized buses.


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CONFIGURATION 1

Loss of one left side generator (network 1):− − Without APU generator available:

− GLC 1 or GLC 3 is open.− ALC is open.− BTC 1 is closed.

− − With APU generator available:− GLC 1 or GLC 3 is open.− ALC is closed.− BTC 1 is closed and BTC 2 is open.

Loss of one right side generator (network 2):− − Without APU generator available:

− GLC 2 or GLC 4 is open.− ALC is open.− BTC 1 is closed.

− − With APU generator available:− GLC 2 or GLC 4 is open.− ALC is closed.− BTC 2 is closed and BTC 1 is open.

Loss of two generators with APU generator available:− GLCs from affected generators are open.− ALC is closed.− BTC 1 and BTC 2 are closed.

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CONFIGURATION 2

Loss of two generators without APU generator available:− GLCs from affected generators are open.− ALC is open.− BTC 1 and BTC 2 are closed.− SBC 1 and SBC 2 are open.

Loss of three generators without APU generator available:− GLCs from affected generators are open.− ALC is open.− BTC 1 and BTC 2 are closed.− SBC 1 and SBC 2 are open.

NOTE: Depending on the amount of load on the remaining buses, anoverload condition may occur. In this case, pilots arerequired to perform an additional load shedding.

Loss of three generators with APU generator available:

− GLCs from affected generators are open− ALC is closed.− BTC 1 and BTC 2 are closed.− SBC 1 and SBC 2 are open.

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CONFIGURATION 3

Loss of all generators:− When the last generator fails, the operational logic configures

the system to dedicate the batteries to supply the EssentialBuses only (electrical emergency condition). In thisconfiguration, the Central DC Bus is also powered to allow theAPU to be started.

− BTC 1, BTC 2, BC 1, SBC 1, SBC 2, BBR 1 and BBR 2 areopen.

− EIC, EBC 1, EBC 2 and BC 2 are closed.

NOTE:- This operational mode is activated for in-flight condition only.- A 1-second time delay is provided to avoid inadvertent

switching to emergency configuration due to electricaltransients.

- If the automatic transfer fails, perform this functionmanually by pressing the Essential Power Button.

- While In-flight, the electrical system is automatically resetif at least one generator is reset and supplying itsassociated bus.

- On the ground, the system can be reset by switching bothBattery Selector Knobs from AUTO to OFF and then backto AUTO.

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CONFIGURATION 4

Short circuit at one DC Bus with all generators on:− Associated battery is removed from affected DC bus through a

fuse.− BTC 1 and BTC 2 are open.− Both GLCs of the affected DC Bus are open, isolating the bus.− Cross-side BTC and EIC are closed and affected side EBC is

energized to maintain both Essential DC Buses energized andbatteries charged.

Short circuit at one DC Bus with loss of one associated generatorand with APU generator:

− Associated battery is removed from the affected DC bus through afuse.

− BTC 1 and BTC 2 are open.− Remaining GLC of the affected DC Bus opens, isolating the bus.− Cross-side BTC and EIC are closed, and affected side EBC is

energized to maintain both Essential DC Buses energized andbatteries charged.

Short circuit at one DC Bus with loss of associated generatorsand with APU generator:

− Both batteries are removed from the affected DC bus throughthe fuses.

− BTC 1 and BTC 2 are open.− EIC closes and EBC of affected side is energized to maintain

the associated Essential DC Bus energized and associatedbattery charged.

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CONFIGURATION 5

Short circuit at one DC Bus with loss of one associated generatorand without APU generator:


− BTC 1 and BTC 2 are open.− Remaining GLC of the affected DC Bus opens, isolating the bus.− Cross-side BTC and EIC close, and EBC of the affected side is

energized to maintain both Essential DC Buses energized andassociated battery charged.

− Both SBCs are open.

Short circuit at one DC Bus with loss of associated generatorsand without APU generator:


− BTC 1 and BTC 2 are open.− EIC closes and EBC of the affected side is energized to

maintain the associated Essential DC Bus energized andassociated battery charged.

− Both SBCs are open.

Short circuit at one DC Bus with loss of associated generators plusone generator of the other side, with or without APU generator:

− The EDL operational sequence is the same as in the previouscondition.

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01 JUNE 20, 2002


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MANUAL

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NORMAL, ABNORMAL AND EMERGENCY OPERATIONDIAGRAMSThe following diagrams present the Electrical System layout whenoperating in normal, abnormal and emergency condition.

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BATT 1 (2) OVTEMP Associated battery temperatureis above 70°C.

WARNING ELEC ESS XFR FAIL Automatic transfer to electricalemergency condition hasfailed.

GEN 1 (2, 3, 4) OVLD Associated generator current isabove 400 A.

GEN 1 (2, 3, 4) OFFBUS

Associated generator isdisconnected from theelectrical network afterengine stabilization due togenerator channel failure orbutton released.

APU GEN OVLD APU generator current isabove 400 A.

CAUTION

APU GEN OFF BUS APU generator isdisconnected from electricalnetwork, due to open ALC,with APU RPM above 95%plus seven seconds. This iscaused by generator channelfailure or button released.

APU CNTOR CLSD APU Starting Contactor (ASC)or Line Contactor (ALC) isinadvertently closed.

DC BUS 1 (2) OFF Associated DC Bus isdeenergized.For airplanes up to S/N 854, ifDC Bus 1 is deenergized theinverter becomes inoperative.

ESS BUS 1 (2) OFF Associated Essential Bus isdeenergized.

SHED BUS 1 (2) OFF Associated Shed Bus isdeenergized.

BATT 1 (2) OFF BUS Associated battery isdisconnected from theelectrical network.


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EICAS MESSAGES (continued)TYPE MESSAGE MEANING

BKUP BATT OFF BUS Backup battery isdisconnected from theelectrical network.

CAUTIONELEC EMERGABNORM

Improper transfer to electricalemergency condition hasoccurred.

115 VAC BUS OFF 115 VAC bus is deenergized(applicable to airplanes up toS/N 854).

ADVISORY GEN 1 (2, 3, 4) BRGFAIL

Associated generator bearinghas failed.

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MANUAL

ELECTRICAL


3 Code

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ELECTRICAL SYSTEM PANEL

AIRPLANES UP TO S/N 854

1 - GENERATOR BUTTON− Connects (pressed) or disconnects (released) the associated

generator to/from the respective DC Bus.− Pressing and depressing the Generator Button causes all GCU

latches protection circuits to be reset if the associated generatoris running.

− A striped bar illuminates inside the button when it is released.

2 - GROUND POWER UNIT BUTTON− Connects (pressed) or disconnects (released) the GPU to/from

the electrical system.− A GPU AVAIL inscription illuminates, in the upper half of the

button, when the GPU is properly connected to the airplanereceptacle and ready to supply power. The GPU AVAILinscription extinguishes when the button is pressed and theexternal power is connected to the electrical network.

− A striped bar illuminates inside the button when it is pressed.

3 - APU STARTER GENERATOR BUTTON− Connects (pressed) or disconnects (released) the APU starter

generator, when APU RPM is above 95%, plus 7 seconds.− A striped bar illuminates inside the button when it is released.

4 - BATTERY SELECTOR KNOBOFF - Respective battery contactor is kept open, disconnecting

the associated battery from the electrical system.AUTO - The actuation of the respective battery contactor is

controlled according to the Electrical Distribution Logic.

5 - ESSENTIAL POWER BUTTON (guarded)− When pressed the system overrides the automatic transfer to

the electrical emergency circuitry, connecting the batteriesdirectly to essential buses, regardless of any other commandfrom the Electrical Distribution Logic.

− When released, the power contactors operate automaticallyaccording to the Electrical Distribution Logic.



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2-05-15 Page

4 Code

01 REVISION 1

6 - SHED BUSES SELECTOR KNOBOVRD - Closes the Shed Buses Contactors, provided the airplane

is on ground and at least one generator is operative.AUTO - Controls the operation of Shed Buses Contactors according

to the Electrical Distribution Logic.OFF - Deenergizes the Shed Buses manually regardless of any

other command from the Electrical Distribution Logic.

7 - AVIONICS MASTER BUTTONS− Connect (pressed) or disconnect (released) the navigation and

communication equipment supplied by the avionics switchedbuses.


8 - BACKUP BATTERY BUTTON− Connects (pressed) or disconnects (released) the backup

battery to/from the electrical system.− A striped bar illuminates inside the button when it is released.

9 - AC POWER BUTTON− Connects (pressed) or disconnects (released) the inverter

to/from the system.− A striped bar illuminates inside the button when it is released.

10 - BUS TIES SELECTOR KNOBOVRD - Bus Tie Contactors (BTCs) are kept closed regardless of

Electrical Distribution Logic, provided that no overcurrent is detected by one of the five GCUs.

AUTO - Controls the operation of the BTCs according to the Electrical Distribution Logic.

OFF - Opens the BTCs and EIC regardless of any other command from the Electrical Distribution Logic.

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


5 Code

01

ELECTRICAL SYSTEM PANEL FOR AIRPLANES UP TO S/N 854


MANUAL

2-05-15 Page

6 Code

01 REVISION 7


AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


7 Code

01

AIRPLANES S/N 863 AND ON

1 - GENERATOR BUTTON− Connects (pressed) or disconnects (released) the associated

generator to/from the respective DC Bus.− Pressing and depressing the Generator Button causes all GCU

latches protection circuits to be reset if the associated generatoris running.


2 - GROUND POWER UNIT BUTTON− Connects (pressed) or disconnects (released) the GPU to/from

the electrical system.− A GPU AVAIL inscription illuminates, in the upper half of the

button, when the GPU is properly connected to the airplanereceptacle and ready to supply power. The GPU AVAILinscription extinguishes when the button is pressed and theexternal power is connected to the electrical network.


3 - APU STARTER GENERATOR BUTTON− Connects (pressed) or disconnects (released) the APU starter

generator, when APU RPM is above 95%, plus 7 seconds.− A striped bar illuminates inside the button when it is released.

4 - BATTERY SELECTOR KNOBOFF - Respective battery contactor is kept open, disconnecting

the associated battery from the electrical system.AUTO - The actuation of the respective battery contactor is

controlled according to the Electrical Distribution Logic.

5 - ESSENTIAL POWER BUTTON (guarded)− When pressed the system overrides the automatic transfer to

the electrical emergency circuitry, connecting the batteriesdirectly to essential buses, regardless of any other commandfrom the Electrical Distribution Logic.

− When released, the power contactors operate automaticallyaccording to the Electrical Distribution Logic.



MANUAL

2-05-15 Page

8 Code

01 REVISION 7

6 - SHED BUSES SELECTOR KNOBOVRD - Closes the Shed Buses Contactors, provided the airplane

is on ground and at least one generator is operative.AUTO - Controls the operation of Shed Buses Contactors according

to the Electrical Distribution Logic.OFF - Deenergizes the Shed Buses manually regardless of any

other command from the Electrical Distribution Logic.

7 - AVIONICS MASTER BUTTONS− Connect (pressed) or disconnect (released) the navigation and

communication equipment supplied by the avionics switchedbuses.


8 - BACKUP BATTERY BUTTON− Connects (pressed) or disconnects (released) the backup

battery to/from the electrical system.− A striped bar illuminates inside the button when it is released.

9 - BUS TIES SELECTOR KNOBOVRD - Bus Tie Contactors (BTCs) are kept closed regardless of

Electrical Distribution Logic, provided that no overcurrent is detected by one of the five GCUs.

AUTO - Controls the operation of the BTCs according to the Electrical Distribution Logic.

OFF - Opens the BTCs and EIC regardless of any other command from the Electrical Distribution Logic.

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


9 Code

01

ELECTRICAL SYSTEM PANEL FOR AIRPLANES S/N 863 AND ON


MANUAL

2-05-15 Page

10 Code

01 REVISION 7

MFD ELECTRICAL PAGE

1 - LABELS AND UNITS− Labels and units are always white.

2 - GENERATOR VOLTAGE AND CURRENT INDICATION

VOLTAGE:− Digits are green and boxes are white during normal operation.− Digits and boxes are amber when the generator is inadvertently

off bus.− Ranges from 0 to 40.0 V, with a resolution of 0.1 V.

CURRENT:− Digits are green and boxes are white during normal operation.− Digits and boxes are amber when the generator is inadvertently

off bus or when the current is higher than 400 A.− Ranges from 0 to 600 A, with a resolution of 5 A.

NOTE: The APU indication is removed when the APU is notavailable and/or the APU Master Selector is set to the OFFposition with APU RPM below 10%.

3 - DC BUS INDICATION− Green when bus is energized.− Amber when bus is off.

4 - GPU VOLTAGE INDICATION− Digits are always green.− Box is always white.− Ranges from 0 to 40.0 V, with resolution of 0.1 V.

NOTE: GPU voltage indication is removed in flight.

5 - BUS LINES INDICATION− Bus lines are always white.

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


11 Code

01

6 - BATTERY VOLTAGE AND TEMPERATURE INDICATION

VOLTAGE:− Digits are green and boxes are white during normal battery

operation.− Digits and boxes are amber when the battery is inadvertently off

bus.− Ranges from 0 to 40.0 V, with a resolution of 0.1 V.

TEMPERATURE:− Boxes are white during battery normal operation.− Boxes are amber when the battery is off bus.− Digits are green when the temperature is below 70°C.− Ranges from –40°C to 150°C, with a resolution of 1°C.− Digits and boxes are red when the temperature is equal to or

greater than 70°C.

NOTE: The red alerts supersede any other condition.

ELECTRICAL PAGE ON MFD


MANUAL

2-05-15 Page

12 Code

01 REVISION 7


AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


1 Code

01

CIRCUIT BREAKER PANEL AND LOADDISTRIBUTION

CIRCUIT BREAKER PANELThe Circuit Breaker Panel is divided in areas associated to electricalsystem buses.

Columns and lines on the circuit breaker panel are identified throughan alphabetic (for the lines) and numeric (for the columns) code.


MANUAL

2-05-20 Page

2 Code

01 REVISION 7


AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


3 Code

01

CIRCUIT BREAKER PANEL MAP FOR AIRPLANES UP TO S/N 854

CIRCUIT BREAKER PANEL MAP FOR AIRPLANES S/N 863 AND ON


MANUAL

2-05-20 Page

4 Code

01 REVISION 7

CIRCUIT BREAKER PANEL FOR AIRPLANES UP TO S/N 854(TYPICAL I)

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


5 Code

01

CIRCUIT BREAKER PANEL FOR AIRPLANES UP TO S/N 854(TYPICAL I)


MANUAL

2-05-20 Page

6 Code

01 REVISION 7

CIRCUIT BREAKER PANEL FOR AIRPLANES UP TO S/N 854(TYPICAL II)

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


7 Code

01

CIRCUIT BREAKER PANEL FOR AIRPLANES UP TO S/N 854(TYPICAL II)


MANUAL

2-05-20 Page

8 Code

01 REVISION 7

CIRCUIT BREAKER PANEL FOR AIRPLANES S/N 863 AND ON

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


9 Code

01

CIRCUIT BREAKER PANEL FOR AIRPLANES S/N 863 AND ON


MANUAL

2-05-20 Page

10 Code

01 REVISION 7

DC BUS LOAD DISTRIBUTION (TYPICAL)The following list identifies the DC buses and the equipment poweredby them. Optional equipment are preceded by an asterisk (*).

DC BUS 1AILERON CONTROL SYSTEM 1AIR/GND POSITION SYSTEM AAOA 1 SENSOR HEATINGBRAKES TEMPERATURE INDICATION OUTBDCABIN LIGHTING 1CENTRAL MAINTENANCE COMPUTERCLEAR ICE DETECTION SYSTEM - CHANNEL 1COCKPIT READING LIGHTCOURTESY/STAIR LIGHTS 2CREW PEDAL ADJUSTMENTCREW SEAT ADJUSTMENT 1EICAS POWER (DAU 1B)ELECTRICAL FLIGHT IDLE STOP 1ELECTRONIC BAY COOLING (EXHAUST 1)ELECTRONIC BAY COOLING (RECIRC 2)EMER/PARKING BRAKEENG 1 FUEL PUMPS 1CENG 1 THRUST REVERSER COMMANDENGINE 1 LIP ANTI-ICEFLAP POWER/COMMAND 1FLOOD/STORM LIGHTSFUEL PRESSURE REFUELING FWD/AFT A1FUEL PRESSURE REFUELING FWD 1FUEL TRANSFER 1GROUND SPOILER OUTBDHYDRAULIC ELECTRIC PUMP 2HYDRAULIC GEN SYS 2 INDICATIONICE DETECTOR 1

∗ INVERTERLANDING LIGHTS 1LAVATORY FLUSHLAVATORY LIGHTSLAVATORY SMOKE DETECTORLAVATORY WATER DRAIN HEATERLOGOTYPE LIGHTSMAIN DOOR CONTROL 1NAVIGATION LIGHTSOVERHEAD PANEL LIGHTINGPACK VALVE 1PASS CABIN LIGHTS 1PASSENGER SIGNSPITCH TRIM 1PITOT 1 HEATINGPNEUMATIC HSV 1PRESSURIZATION CONTROLSPEED BRAKESTABILISER ANTI-ICE INDICATIONSTATIC PORT HEATING 1STROBE LIGHTSTAT 1 SENSOR HEATING

∗ TCAS 2000WINDSHIELD HEATING 1

∗ WINDSHIELD WIPER SYSTEM 1WING ANTI-ICE SYSTEMYAW TRIM SYSTEM

DC BUS 2ADC 2 POWER/CONTROL

∗ AHRS 2 POWERAILERON CONTROL SYSTEM 2AIR/GND POSITION SYSTEM CAOA 2 SENSOR HEATINGAURAL WARNING SYSTEM 2BAGGAGE SMOKE DETECTORBRAKES TEMPERATURE INDICATION INBDCABIN RECIRCULATIONCLEAR ICE DETECTION SYSTEM - CHANNEL 2COMPARTMENT LIGHTSCOPILOT'S CLOCKCREW SEAT ADJUSTMENT 2DEFUELINGDISPLAY PRCS/CONTROL POWER 2 (IC2)EICAS POWER (DAU 2B)ELECTRICAL FLIGHT IDLE STOP 2ELECTROMECHANICAL GUST LOCKELECTRONIC BAY COOLING (RECIRC 1)ELECTRONIC BAY COOLING (EXHAUST 2)ENG 2 FUEL PUMPS 2CENG 2 THRUST REVERSER COMMANDENGINE 2 LIP ANTI-ICEENGINE VIBRATION SENSORSFLAP POWER/COMMAND 2FUEL FUSELAGE PUMPS AFT/FWD 2AFUEL FUSELAGE PUMPS FWD 2BFUEL TRANSFER 2GASPER FANGROUND SPOILER INBDHYDR ELECTRIC PUMP 1HYDR GEN SYS 1 INDICATIONICE DETECTOR 2INSPECTION LIGHTSIRS 2 POWERLANDING GEAR DOOR COMMANDLANDING LIGHTSOBSERVER AUDIO (INTPH 3)OVERHEAD PANEL LIGHTINGPACK VALVE 2PASSENGER CABIN LIGHTS 2/3PITOT 2 HEATINGPNEUMATIC HSV 2RED BEACON LIGHTSROLL TRIM SYSTEMSENSORS HEATING CONTROLSPOILER INDICATIONSPS (SHAKER 2/CHANNEL 2)SPS PUSHERSTABILIZER ANTI-ICE SYSTEMSTATIC PORT HEATING 2STEERING SYSTEMTAT 2 SENSOR HEATINGWINDSHIELD WIPER SYSTEM 2WING ANTI-ICE 1 INDICATION

AIRPLANEOPERATIONS

MANUAL

ELECTRICAL


11 Code

01

AVIONIC SWITCHED DC BUS 1AAUTOPILOT 1DME 1HF 1MFD 2 POWER

∗ MLS 1 POWER/CONTROLPFD 1 POWERRADIO ALTIMETER 1

AVIONIC SWITCHED DC BUS 1B∗ CMU MARK III∗ FLITEFONE∗ FMS SYSTEM 1 DATA LOADER∗ FMS SYSTEM 1 COMPUTER∗ FMS SYSTEM 1 CDU

RADAR SYSTEM∗ TDR 1 POWER/CONTROL∗ VHF SYSTEM 3

SHED DC BUS 1COCKPIT RECIRCULATION

COCKPIT IFE & PC POWER PANEL ENTERTAINMENT CABINET CB PANEL

GALLEY OVEN POWERNOSE LANDING LIGHTS

∗ MUSIC PC POWER INVERTER∗ PRE RECORD ANNOUNCEMENTS (PRA)

READING LIGHTS/ATTENDANT CALL 1 WATER HEATING INVERTER 1∗ SELCAL SYSTEM

HOT BUS 1EMERGENCY LOCATOR TRANSMITER (ELT)ENG 1 FIRE EXTINGUISHING (BTL A1)ENG 2 FIRE EXTINGUISHING (BTL A2)FUEL PRESSURE REFUELING 3FUEL SHUTOFF VALVES 1HYDRAULIC SHUTOFF VALVE 1

BACKUP ESSENTIAL BUS∗ AHRS 1 POWER

DATA ACQUISITION UNIT ½DISPLAY PRCS/CONTROL POWER 1EICAS POWERIRS 1 POWER

(#) Applicable only if DUAL FMS is installed

AVIONIC SWITCHED DC BUS 2AAUTOPILOT 2DME 2

∗ FMS SYSTEM 2 DATA LOADER (#)∗ FMS SYSTEM 2 COMPUTER (#)∗ FMS SYSTEM 2 CDU (#)

MFD 1 POWER∗ MLS 2 POWER/CONTROL

PFD 2 POWER∗ RADIO ALTIMETER 2

TUNING BACKUP CONTROL HEADVHF SYSTEM 2

AVIONIC SWITCHED DC BUS 2B∗ ADF 2∗ GPS

HF 2∗ OMEGA

TDR 2 POWER/CONTROLVOR/ILS/MB 2

SHED DC BUS 2AIR COMPRESSORCABIN RECIRCULATIONCONVECTION OVENFLASHLIGHT

∗ GALLEY∗ GALLEY COFFEE MAKER POWER

MICROWAVE OVENREADING LIGHTS/ATTENDANT CALL 2/3TAXI LIGHTSWATER HEATING INVERTER 2WINDSHIELD HEATING 2

HOT BUS 2COURTESY/STAIR LIGHTS 1ENG 1 FIRE EXTINGUISHING (BTL B 1)ENG 2 FIRE EXTINGUISHING (BTL B 2)FUEL SHUTOFF VALVES 2FUSELAGE FUEL ISOLATION VALVEFUSELAGE FUEL VENT VALVEHYDRAULIC SHUTOFF VALVE 2MAIN DOOR CONTROL 2

BACKUP HOT BUSAPU GENERATIONDC DISTRIBUTIONDC GENERATION 1DC GENERATION 2DC GENERATION 3DC GENERATION 4ISIS


MANUAL

2-05-20 Page

12 Code

01 REVISION 7

BACKUP BUS 1NONE

ESSENTIAL DC BUS 1ADC 1 POWER/CONTROL

∗ AHRS 1 POWERAIR/GND POSITION SYSTEM BAPU BLEEDAURAL WARNING SYSTEM 1COCKPIT DOME LIGHTSDISPLAY PRCS/CONTROL POWER 1 (IC 1)EICAS DISPLAY POWEREICAS POWER (DAU 1A)ENG 1 FIRE DETECTION 1ENG 1 FUEL PUMPS 1AENG 2 FUEL PUMPS 2BENGINE 1 FADEC A POWERENGINE 2 FADEC A POWERENGINE 1 STARTINGENGINES N2 SIGNALS 1AENGINES N2 SIGNALS 2AFDR MANAGEMENTFUEL QUANTITY INDICATION 1LANDING GEAR CONTROL (DOWN OVRD)LANDING GEAR NOSE INDICATION 1IRS POWER 1PASSENGER OXYGEN SYSTEM 1PILOT/COPILOT AUDIO SYSTEM (INTPH 1)PILOT'S CLOCKPILOT'S PANEL LIGHTINGPNEUMATIC 1 (EBV 1)RAM AIR DISTRIBUTIONRMU 1 POWER/CONTROLRUDDER CONTROL SYSTEM 2SPS (SHAKER 1/CHANNEL 1)VHF SYSTEM 1

AVIONIC SWITCHED ESSENTIALDC BUS 1

ADF 1VOR/ILS/MB 1

BACKUP BUS 2∗ AHRS 2 POWER

IRS 2 POWER

ESSENTIAL DC BUS 2AIR/GND POSITION SYSTEM DAPU CONTROLAPU FIRE DETECTIONAPU FIRE EXTINGUISHINGAPU FUEL FEEDCOPILOT'S PANEL LIGHTINGCROSSBLEEDEICAS POWER (DAU 2A)EMERGENCY LIGHTING CONTROLENG 2 FIRE DETECTION 2ENG 1 FUEL PUMPS 1BENG 2 FUEL PUMPS 2AENGINE 1 FADEC B POWERENGINE 2 ANTI-ICE INDICATIONENGINE 2 FADEC B POWERENGINE 2 STARTINGENGINES N2 SIGNALS 1BENGINES N2 SIGNALS 2BFUEL CROSSFEEDFUEL QUANTITY INDICATION 2ISIS (AIRPLANES S/N 484, 495, 528, 540, 555)LANDING GEAR BRAKES INBOARDLANDING GEAR CONTROLLANDING GEAR NOSE INDICATION 2PASSENGER OXYGEN SYSTEM 2PEDESTAL PANEL LIGHTINGPILOT/COPILOT AUDIO SYSTEM (INTPH 2)PITCH TRIM 2PITOT HEATING 3PNEUMATIC 2 (EBV 2)PUBLIC ADRESSRMU 2 POWER/CONTROLRUDDER CONTROL SYSTEM 1STANDBY ALTIMETERSTANDBY ATTITUDE INDICATORVOICE RECORDERWING ANTI-ICE INDICATION 2

AVIONIC SWITCHED ESSENTIALDC BUS 2

NONE

AIRPLANEOPERATIONS

MANUAL

EQUIPMENTAND FURNISHINGS


1 Code

01

SECTION 2-02

EQUIPMENT AND FURNISHINGSTABLE OF CONTENTS

Block Page

Cockpit ............................................................................... 2-02-05 ..01Pilot Seats ...................................................................... 2-02-05 ..01Pilot Seat Controls.......................................................... 2-02-05 ..02Pilot Seat Adjustment ..................................................... 2-02-05 ..04Pedal Adjustment ........................................................... 2-02-05 ..05Observer Seat ................................................................ 2-02-05 ..06Direct Vision Windows.................................................... 2-02-05 ..08

Attendant Station and Seat (Optional) ............................... 2-02-10 ..01

Galley ................................................................................. 2-02-15 ..01Forward Galley - Main Components............................... 2-02-15 ..02Galley Electrical System................................................. 2-02-15 ..06Galley Components ........................................................ 2-02-15 ..08Galley Lighting ................................................................ 2-02-15 ..10Water.............................................................................. 2-02-15 ..12Controls and Indicators................................................... 2-02-15 ..13

Passenger Seats................................................................ 2-02-20 ..01Controls and Indicators................................................... 2-02-20 ..04

Escutcheons....................................................................... 2-02-25 ..01Controls and Indicators................................................... 2-02-25 ..02

Closets ............................................................................... 2-02-30 ..01Components ................................................................... 2-02-30 ..02Controls and Indicators................................................... 2-02-30 ..14

Partitions ............................................................................ 2-02-35 ..01Components ................................................................... 2-02-35 ..04Controls and Indicators................................................... 2-02-35 ..14

Water and Waste ............................................................... 2-02-40 ..01Water.............................................................................. 2-02-40 ..01Waste Disposal .............................................................. 2-02-40 ..06

Airstair Main Door .............................................................. 2-02-45 ..01EICAS Message ............................................................. 2-02-45 ..01Controls and Indicators................................................... 2-02-45 ..02Main Door Acoustic Curtain............................................ 2-02-45 ..06


AIRPLANEOPERATIONS

MANUAL

2-02-00 Page

2 Code

01 REVISION 4

Access Doors and Hatches ................................................2-02-50.. 01Baggage Door.................................................................2-02-50.. 04Compartment Hatches....................................................2-02-50.. 06Refueling Panel Access Door .........................................2-02-50.. 07Lavatory/Baggage Compartment Access Door ..............2-02-50.. 07Emergency Exit Hatches ................................................2-02-50.. 08Doors and Hatches Indication on MFD...........................2-02-50.. 08Forward Lavatory Doors .................................................2-02-50.. 10

AFT Lavatory.....................................................................2-02-55.. 01Vanity Assembly .............................................................2-02-55.. 04Toilet Section Components.............................................2-02-55.. 04Lavatory Electrical Installation ........................................2-02-55.. 06Controls and Indicators...................................................2-02-55.. 08

Forward Lavatory................................................................2-02-57.. 01Toilet Section Components.............................................2-02-57.. 04Lavatory Electrical Installation ........................................2-02-57.. 06Controls and Indicators...................................................2-02-57.. 06

Pilot and Passenger Convenience Items ...........................2-02-60.. 01

PC Power System ..............................................................2-02-65.. 01Ground Fault Circuit Interrupter ......................................2-02-65.. 02Controls and Indicators...................................................2-02-65.. 03

In-Flight Entertainment System ..........................................2-02-70.. 01Main Components...........................................................2-02-70.. 02Controls and Indicators...................................................2-02-70.. 04

SATCOM System ...............................................................2-02-75.. 01

Telephone System (Optional).............................................2-02-80.. 01

AIRPLANEOPERATIONS

MANUAL



1 Code

01

COCKPIT

PILOT SEATS

The pilot seats are fixed to slide rails which permit fore and aftadjustments. When the seats are in their aft most position, a lateralmovement is also available, in order to facilitate crew access to theseat. The seats are fitted with adjustable armrest, seat backs, thighsupport and lumbar position, and can also be adjusted for height.Backrest inclination, thigh support and lumbar positions are hydro-mechanically adjusted. The armrest adjustment, and seat fore, aft andlateral adjustments are made mechanically.

The pilot and copilot seats are identical, except for the symmetricalarrangement of the controls. Controls on the pilot’s seat are on theopposite side from those on the copilot’s seat.

A switch installed in the seat allows height adjustment, which isperformed by an electrical actuator. In case of electrical actuatormalfunction height adjustment may also be accomplished manually byattaching a crank to the actuator and rotating it. Extension or retractionof the actuator rod connected to the seat structure permits verticaldisplacement.

The crew seat belts consist of five straps. The left (for the pilot seat)and right (for the copilot seat) lap belt straps are permanently fixed to arotary buckle, provided with quick-release latch locks that are operatedby turning the existing rotary device on the buckle face. The two upperstraps are connected to an inertia reel attached to the seat backrest,which allows the pilot to bend forward in normal, slow movements.Abrupt movements or high acceleration locks the upper straps,preventing the pilot from impacting against the instrument panel. Theinertia reel can be mechanically locked through a lever installed on theseat.


AIRPLANEOPERATIONS

MANUAL

2-02-05 Page

2 Code


PILOT SEAT CONTROLS

1 - SEAT FORE/AFT AND LATERAL ADJUSTMENT LEVER− Pulling the lever up, the seat is free to slide along its rails.

Lateral movement is allowed only when the seat is at the aftstop.

− Releasing the lever, the seat is locked. Fore/aft movement haspredetermined fixed positions. Lateral movement has only theleft and right stops.

2 - SEAT HEIGHT ADJUSTMENT BUTTON (spring loaded, centeroff rocker button)− Pressing the button up or down causes the seat to raise or to

lower respectively, provided the airplane is energized.

3 - BACKREST INCLINATION ADJUSTMENT BUTTON− Pressing the button allows the occupant to select the required

inclination by pressure exerted upon the backrest.− Releasing the button, backrest is retained in the desired

position.

4 - LUMBAR ADJUSTMENT WHEEL− When rotated, provides in and out lumbar adjustment.

5 - THIGH SUPPORT ADJUSTMENT WHEEL− When rotated, provides thigh support height adjustment.

6 - ARMREST ANGLE ADJUSTMENT WHEEL− When rotated, allows armrest adjustment to the desired angle.

7 - INERTIA REEL LOCK LEVERLOCK - Locks the inertia reel in the current position.UNLOCK - Unlocks the inertia reel, permitting normal belt

movement.

8 - HEIGHT ADJUSTMENT LEVER BACK-UP− When attached to the height adjustment actuator and rotated, it

causes the seat to raise or to lower.

AIRPLANEOPERATIONS

MANUAL



3 Code

01

PILOT SEAT CONTROLS


AIRPLANEOPERATIONS

MANUAL

2-02-05 Page

4 Code


PILOT SEAT ADJUSTMENT

Seat adjustment should be accomplished to accommodate the pilot’seye level and position best suited for control column actuation. Theseat should be moved up or down until the pilot’s line of sight reachesthe same horizontal plane of a sight device made up of two whitespheres and a black sphere. Then, move the seat fore or aft until theopposite white sphere is aligned with the black one. The seat shouldnot be moved anymore. To adjust the rudder pedals, refer to PEDALADJUSTMENT.

PILOT SEAT ADJUSTMENT

AIRPLANEOPERATIONS

MANUAL



5 Code

01

PEDAL ADJUSTMENT

Toggle switches installed on the pilot and copilot’s panels allowsrudder pedals adjustment, which is performed by electric actuators.Setting the switch up or down signals the actuator to move the pedalsfore or aft, to assure the pilot’s comfort and a full rudder throw from theadjusted seat position.

PEDAL ADJUSTMENT SWITCH


AIRPLANEOPERATIONS

MANUAL

2-02-05 Page

6 Code


OBSERVER SEAT

The observer seat is installed behind the co-pilot seat. When in use, itlies in front of the cockpit door, and when not in use, it folds up androtates away from the door area, stowing against the right side of thecockpit partition.

The cockpit door can be opened or closed either with the observer seatin use or stowed.

OBSERVER SEAT

AIRPLANEOPERATIONS

MANUAL



7 Code

01

OBSERVER SEAT


AIRPLANEOPERATIONS

MANUAL

2-02-05 Page

8 Code


DIRECT VISION WINDOWS

The normal position for the direct vision windows is closed. However,they may be partially opened on the ground, and may be totallyremoved in case of loss of visibility through the windshield or forcockpit emergency evacuation. Placing respective pilot seat to the aftmost position facilitates window removal.

A yellow pin protrudes near the opening handle when the window is notproperly locked in the closed position, indicating the unlockedcondition.

A WINDOW NOT CLOSED inscription on the window front frame willbe visible when the window is not properly closed.

DIRECT VISION WINDOW REMOVAL

AIRPLANEOPERATIONS

MANUAL


JUNE 20, 2002 2-02-10 Page

1 Code

01

ATTENDANT STATION AND SEAT

The flight attendant station is positioned at the cockpit partition, closeto the main door. This fold-away type seat avoids interference with thedoor passage way.

FLIGHT ATTENDANT STATION (OPTIONAL)


AIRPLANEOPERATIONS

MANUAL

2-02-10 Page

2 Code

01 JUNE 20, 2002


AIRPLANEOPERATIONS

MANUAL



1 Code

01

GALLEY

A full height forward galley, matching the cabin´s contours, provides acentral location to perform food and beverage preparation.

The galley is installed in the forward section of the passenger cabinarea and includes as standard items an oven, a microwave oven,coffee maker, trash container (13 gal / 49.2 liters volume) drawer withfire extinguisher, control panels for the water system, emergencylights, oven and galley, pop-out auxiliary work surface, paper towel rollholder, cutting boards, storage provisions for utensils and napkins,flatware storage, ice compartment drawers, glassware storage, crystaland chinaware storage and carrier, seasoning and spices storage, sinkwith hot water, bottle and can storage, juice can storage, liquordecanters or miniature storage, work table, work light, electrical circuitbreakers panels, miscellaneous storage and in line water heaters andfilters.

The forward galley is composed of an aft cabinet assembly, an uppercabinet assembly, a lower cabinet assembly, and a close-out panelassembly. The cabinet structures are made of lightweight honeycombpanel and its exterior is covered with a decorative finish.

The forward galley provides an area to house two oxygen cylinders forpassenger main oxygen supply.

The forward galley assembly is provided with an electrical installation,a plumbing installation, an electrical hot water heater and heatedoverboard drain system.

The electrical installation is provided with a galley control module andelectrical installation hardware. The galley control module controls thegalley lights, valance lights, wash lights and passenger cabintemperature control.

The plumbing installation allows drainage of liquids from thecountertop. The waste compartment is provided for the stowage offood waste. The waste compartment comprises a trash container withtrash bag and fire extinguisher.

There is a water system that stores and supplies drinking water for useby cabin occupants and crew members, and both to the galley andlavatory washbasins.

A pullout table assembly is also provided to give extra room for foodand beverage preparation.


AIRPLANEOPERATIONS

MANUAL

2-02-15 Page

2 Code

01 REVISION 4

FORWARD GALLEY - MAIN COMPONENTS

FOR AIRPLANES UP TO S/N 685 (INCLUSIVE).The following list enumerates the main components found in theforward galley, the figure provided on the next page indicating thelocation of these components.

1 - GLASS STORAGE RACK COMPARTMENT2 - VACUUM CARAFE SETS3 - COFFEE MAKER4 - MISCELLANEOUS COMPARTMENTS5 - ROLL DOOR6 - OVEN CONTROL PANEL7 - GALLEY CONTROL PANEL8 - MICROWAVE OVEN9 - PULL-OUT TABLE ASSEMBLY10 - WASTE/AUTO FIRE EXTINGUISHING COMPARTMENT11 - ELECTRIC OVEN12 - BOTTLES DRAWER13 - BUS BIN STORAGE14 - SERVING TRAYS/FLATWARE STORAGE15 - CE DRAWER16 - FAUCET ASSEMBLY17 - OXYGEN CYLINDERS18 - PAPER TOWEL HOLDERS

AIRPLANEOPERATIONS

MANUAL



3 Code

01

FORWARD GALLEY(FOR AIRPLANES UP TO S/N 685 (INCLUSIVE))

MAIN COMPONENTS LOCATION


AIRPLANEOPERATIONS

MANUAL

2-02-15 Page

4 Code

01 REVISION 4

FOR AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENTAIRFRAMES.The following list presents the main components on the forward galley,and the figure provided on the next page shows the location of thesecomponents. 1 - CRYSTAL STORAGE COMPARTMENTS2 - MISCELLANEOUS STORAGE COMPARTMENTS3 - EMERGENCY LIGHTS CONTROL PANEL4 - WATER WASTE DRAIN LINES HEATER TEST CONTROL

PANEL5 - GALLEY CONTROL PANEL6 - DRINK BOTTLES AND SEASONING STORAGE

COMPARTMENTS7 - COFFEE MAKER8 - FLATWARE STORAGE COMPARTMENT9 - CLEAN ICE DRAWER10 - LONG NECK STORAGE COMPARTMENT11 - WASTE COMPARTMENT12 - SODA STORAGE COMPARTMENTS13 - MISCELLANEOUS STORAGE COMPARTMENTS14 - CHINAWARE STORAGE COMPARTMENTS15 - WINE STORAGE COMPARTMENT16 - ELECTRIC OVEN17 - MICROWAVE OVEN18 - CRYSTAL STORAGE COMPARTMENT19 - FOOD TRAYS STORAGE COMPARTMENT20 - MISCELLANEOUS STORAGE COMPARTMENTS21 - OXYGEN CYLINDER22 - FAUCET ASSEMBLY23 - MUG STORAGE COMPARTMENT24 - ROLL-UP DOOR25 - OVEN CONTROL PANEL

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FORWARD GALLEY (FOR AIRPLANES FROM S/N 686(INCLUSIVE) AND SUBSEQUENT AIRFRAMES)

MAIN COMPONENTS LOCATION


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GALLEY ELECTRICAL SYSTEM

The galley´s electrical installation supplies 28 V DC power to the galleyfor distribution to the associated components.

The galley electrical installation receives 28 V DC from the RH powercontrol and distribution box, located on the aft console, and supplieselectrical power to the main galley´s components. The galley´selectrical installation also receives 28 V DC from the entertainmentcabinet circuit breaker panel, located in the entertainment cabinet. Thispowers the water heater, 60 Hz outlet GFI (Ground Fault Interrupter),credenza lights and galley lights.

The electric oven, microwave oven and coffee maker are installed inan enclosure with a door. The oven and the microwave oven have amicroswitch installed on the door that controls a relay (located inside ofeach component) to cut their power supply when the door is closed.

The coffee maker has a microswitch installed at the door that controlsa relay to cut off the power supply from it when the door is closed.

The galley´s electrical installation is also provided with a galley controlpanel, galley lighting, advisory signs and a fuse box. The fuse box,installed behind the galley´s structure, protects the electric oven,microwave oven and coffee maker. The galley also provides a 60 Hzoutlet GFI, and electrical connectors to connect the galley components.

The galley´s electrical installation includes the following components:

− Microwave oven− Electric oven− Oven controller− Coffee maker− Water heater− Galley control panel− Galley lighting

The location of these components is presented in the figure on thenext page.

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GALLEY ELECTRICAL INSTALLATIONMAIN COMPONENTS LOCATION


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GALLEY COMPONENTS

MICROWAVE OVEN

Microwave cooking time is set by the controls on the front of the unit. Itcontrols the oven operation and provides an audible signal when thecooking cycle is complete.

The microwave oven door operates a safety interlock switch, whichdisables the cooking cycle while the door is open or left ajar. There isan additional toggle-type latch installed at the top of the door to furtherensure that the door will remain closed during the cooking cycle.

The microwave oven is equipped with a pressure switch that cuts offthe microwave oven power supply when the passenger cabin pressureexceeds 10000 ft. The microwave oven power supply will be restoredwhen the pressure falls below 9000 ft. The solid state timer of the ovenis not affected by momentary power interruptions.

ELECTRIC OVEN

The electric oven has a temperature range from 150°F to 450°F(65.5°C to 232.2°C) with overheat protection. A fan circulates air insidethe oven for even heat distribution of ± 10°F (± 12.2°C).

An external oven controller controls the electric oven.

COFFEE MAKER

The coffee maker uses standard coffee filters and regular or drip grindcoffee.

The brewer basket is locked into the coffee maker and can beremoved by lifting up its respective release button. The thermal carafeis locked into the coffee maker and can be removed by pressing itsrespective release button. The coffee maker has automatic fill and isdesigned for operation with pressurized water systems. Hot and coldwater can be dispensed through a spigot on the coffee maker.

The coffee maker has redundant overheat protection. The watersupply must be turned off and the brewer drain valve on the right handside in the rear of the carafe must be open.

Super fine grinds such as “express” can cause the carafe to overfill.When electrical power is available, a POWER ON red light indicatorwill glow. The BREW button is used to start heating the water. Theflashing green light indicates that the water has not reached the proper

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temperature. The green HOT WATER light will glow to indicate that thewater has reached the proper temperature and the brew process willbegin. The entire brewing cycles takes approximately 6 minutes. Thegreen BREW light turns off to indicate that the brewing process iscompleted.

Press and hold the HOT WATER button. The hot water will come outof the spigot. Press and hold the COLD WATER button. The coldwater will come out of the spigot.

WATER HEATER

The water heater comes on automatically when 28 V DC is applied bythe water heater circuit breaker (located in the entertainment cabinetcircuit breaker panel) closed. The heater maximum temperature is pre-selected to 115°F (46.1°C). It has a capacity of 1.4 liters. The waterheater can not provide continuous hot water supply. Its recovery time isapproximately 15 minutes.

The water heater is protected against overtemperature by anovertemperature switch with manual reset and against overpressure byan overpressure switch.

OVEN CONTROLLER

The oven controller time allows for one hour of cooking time.

A red light will come on to indicate that the oven is on (POWER ON).An amber light will come on to indicate that the heaters are on. Theamber light will start flashing when the oven reaches the selectedtemperature.


The galley control panel houses switches that control area for thepassenger cabin temperature control switches, galley work light switch,forward upwash lights switch, aft upwash lights switch, lavatory callindicator switch, credenza lights switch, galley accent lights switch,galley area lights switch, forward downwash lights switch and aftdownwash lights switch.


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GALLEY LIGHTING

Galley lighting consists of a work light installed above the sink area,two accent lights installed in the glass storage rack compartment andthree area lights installed in the ceiling above the galley for aisleillumination in front of the galley.

The figure on the next page presents the location of galley lightingcomponents.

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GALLEY LIGHTS


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WATER

The water subsystem includes components and equipment to storeand supply drinking water for cabin occupants and crewmembers. Itstores water in a pressurized tank and supplies this water to thewashbasins in the lavatory and forward galley. In some airplanes thewater is stored in two tanks: one in the forward galley and the other inthe lavatory. If the airplane is equipped with a forward lavatory, there isanother water tank for the washbasin.

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FOR AIRPLANES UP TO S/N 685 (INCLUSIVE).


The galley control panel has the ability to control the following items:

1 - CABIN TEMPERATURE− The cabin temperature is displayed in bargraph format. The

cabin temperature can be raised or lowered with the “Up” or“Down” TEMP buttons.

2 - GALLEY WORK LIGHTS− When alternately pressed, turns ON or OFF the galley work

lights.

3 - CABIN WASH LIGHTS− There are controls to select the FWD Upwash, FWD

Downwash, AFT Upwash and AFT Downwash lights. Theselection is DIM/BRIGHT/OFF.

4 - GALLEY AREA LIGHTS− When alternately pressed, turns ON or OFF the galley area

lights.

5 - GALLEY ACCENT LIGHTS− When alternately pressed, turns ON or OFF the galley accent

lights.

6 - CREDENZA LIGHTS− When alternately pressed, turns ON or OFF the credenza lights.

7 - LAVATORY CALL INDICATION− When the LAV CALL button is illuminated, it indicates a

passenger call in the lavatory.

EMERGENCY LIGHTS CONTROL PANEL

1 - EMERGENCY LIGHTS− When alternately pressed, turns ON or OFF the emergency

lights.


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WATER CONTROL PANEL

The water control panel is composed of the following items:

1 - TANK WATER LEVEL INDICATOR− Indicates existing water level inside the storage tank.

2 - COMPRESSOR RESET SWITCH− The air compressor will remain off until the tank pressure drops

down to or below 20 ± 5 psig (the overpressure switch closes)and the compressor resetting button is pressed.

3 - PUSH TO TEST BUTTON− Verifies if the indications in the water system control panel

illuminate.

4 - TANK OVER PRESSURE INDICATOR− Provides indication of the tank overpressure condition. If the

tank pressure rises to 48 ± 2 psig, the tank overpressure switchopens and removes power from the air compressor.

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GALLEY, WATER AND EMERGENCY LIGHTS CONTROL PANELS(FOR AIRPLANES UP TO S/N 685 (INCLUSIVE))


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FOR AIRPLANES FROM S/N 686 (INCLUSIVE) AND ON.

DRAIN HEATER CONTROL PANEL

The drain heater control panel is composed of the following items:

1 - FWD HEATER TEST− Tests the continuity of the forward drain heater.

2 - FWD DRN HEATER− Turns on the forward drain/hose heater.

3 - AFT DRN HEATER− Turns on the aft drain/hose heater.

4 - AFT HEATER TEST− Tests the continuity of the aft drain heater.

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WATER CONTROL PANEL

The water control panel is composed of the following items:

5 - WATER TANK LEVEL INDICATOR− Indicates the existing water level inside the water storage tank.

6 - GALLEY WORK LIGHTS− Turns the galley work lights on/off.

7 - GALLEY AREA LIGHTS− Turns the galley area lights on/off.

8 - FWD UPWASH LIGHTS− Selects the forward upwash lights bright/dimmed (DIM)/off.

9 - AFT UPWASH LIGHTS− Selects the aft upwash lights bright/dimmed (DIM)/off.

10 - AFT DNWASH LIGHTS− Selects the aft downwash lights bright/dimmed (DIM)/off.

11 - FWD DNWASH LIGHTS− Selects the forward downwash lights bright/dimmed (DIM)/off.

12 - LAV WATER ON/OFF− Selects the check carried out by the water system control unit in

the lavatory tank.

13 - GALLEY WATER ON/OFF− Selects the check carried out by the water system control unit in

the galley tank.

14 - LAV WATER LEVEL− Selects the water level indication of the lavatory tank.

15 - GALLEY WATER LEVEL− Selects the water level indication of the galley tank.

16 - ATTNDNT CALL− Clears the attendant call indication.


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EMERGENCY LIGHTS CONTROL PANEL

The emergency lights control panel is composed of the following item:

17 - EMERGENCY LIGHTS SWITCH− When alternately pressed, turns ON or OFF the emergency

lights.− Amber: Indicates that the emergency lights are in normal flight

position.− Green: Indicates that the emergency lights are turned on.

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GALLEY, DRAIN HEATER AND EMERGENCY LIGHTSCONTROL PANELS

(FOR AIRPLANES FROM S/N 686 (INCLUSIVE) ANDSUBSEQUENT AIRFRAMES)


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PASSENGER SEATSPassenger seats have a single control lever for base tracking, hiddenheadrests, color coordinated seat belts webbing, upholstered in leatherper customer specification, thigh rest, lumbar support adjustment,armrests, leg and footrest.

Each Executive Single Seat has forward, aft, swivel and lateralmovement, limited recline control on inside of inboard arm, footrest, lifevest storage in each seat base compartment and escape lighting. Thetwo forward single seats facing the forward bulkhead have berthingcapability.

The Executive Divan (3 place) has three color coordinated seat belts,berthing capabilities, one color coordinated sleeper belt, storage for lifevests and passengers rafts included.

Each Single or Double Executive Seat (except the inboard club seatswhen mentioned below) has a control unit in the sideledge whichallows the management of the reading/table lights (except the inboardclub seats) and have an integrated In-Flight Entertainment Systemmanagement, also in the sideledge, consisting of a headphone jackwith the volume control (all seats), audio (all seats) and video sourceselection (except the inboard club seats).

Each seat has a 6.5” plug-in type LCD monitor, except the inboard clubseats and the divan, that can display the three standard video sources(video cassette player, DVD/CD player and Airshow 400). All seatshave the capability to switch between all video/audio sources listedabove except the inboard club seats. The audio amplifiers supply theaudio to eight (8) speakers located above the tables throughout thecabin. Two subwoofers are installed within the cabin. Each seat hasthe ability to select the audio sources and control the volume for theirassociated headphone jack.

There are two VIP seats in the passenger cabin. The forward VIP seatis the left single seat facing the forward bulkhead. The aft VIP seat(optional) is the left single seat facing the lavatory bulkhead. Theforward VIP seat has the additional capability to control the audiosource and volume for the forward cabin, overhead speaker system.Similarly, the aft VIP seat has the capability to control the audio sourceand volume for the aft cabin, overhead speaker system.

The forward seat on the right side, facing the forward bulkhead, andaside to the VIP seat, is also provided with a SVGA port to displayMicrosoft PowerPoint presentations on the optional forward 15.1”


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bulkhead monitor located on the also optional closet placed in the leftside of the aircraft. Also, a serial printer port connection, located in theforward seat on the right side, facing the forward bulkhead, allows datato be transferred to an optional serial printer/fax in the credenza.

An air-to-ground telephone system is installed with two handsets. Onehandset is located at the forward VIP seat and another at the mid-section seating group, in the conference table.

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PASSENGER SEATS


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PASSENGER CONTROL PANELS

The passenger control panels are installed on the sideledge for allseats except the VIP seat, or in the inboard club seats. These panelshave the ability to control the following items:

1 - READING LIGHT− The reading light can be turned ON/OFF by pressing the “READ

LIGHT” ON/OFF switch (green = ON, amber = OFF) on thepassenger control panel at each seat.

2 - TABLE LIGHT− The table light can be turned ON/OFF by pressing the “TABLE

LIGHT” ON/OFF switch (green = ON, amber = OFF) on thepassenger control panel at each seat.

3 - PERSONAL VIDEO SELECTION− In the sideledge, the “PERSONAL VIDEO SELECT” button can

be used to select “VCR”, “DVD1” or “DVD2”, which will bedisplayed on the LED display. The VCR/DVD video will appearclear and sharp on the individual 6.5” monitor installed at theseat location.

− In the inboard club seats, the “UP” and “DOWN” buttons can beused to select “VCR”, “DVD1” or “DVD2”, which will be displayedon the LED display. The VCR/DVD video will appear clear andsharp on the individual 6.5” monitor installed at the seat location.

4 - HEADSET AUDIO SELECTION− Using the “HEADSET AUDIO SELECT” switch, change the

audio until the VCR, DVD1 or DVD2 audio can be heard fromheadset. The audio source selected will be displayed on theLED display.

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DOUBLE EXECUTIVE INBOARD CLUB SEATSPASSENGER CONTROL PANELS


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SINGLE EXECUTIVE SEATS AND DOUBLE EXECUTIVE SEATSSIDELEDGE PASSENGER CONTROL PANEL

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VIP CONTROL PANELS

FORWARD VIP CONTROL PANEL

The forward VIP control panel is located in the left-hand side panelnear the forward VIP seat. There is a headset jack andvolume/temperature “UP” and “DOWN” control buttons. The audio willcorrespond to the chosen video selection, however, it can be selectedseparately from the video. This panel has the ability to control thefollowing items:

1 - FORWARD UPWASH LIGHTS− The forward cabin upwash lights can be turned ON or OFF by

pressing the “FWD UPWASH LIGHTS” ON/OFF switch (green =ON, amber = OFF).

2 - AFT UPWASH LIGHTS− The aft cabin upwash lights can be turned ON or OFF by

pressing the “AFT UPWASH LIGHTS” ON/OFF switch (green =ON, amber = OFF).

3 - READING LIGHT− The reading light can be turned ON or OFF by pressing the

“READ LIGHT” ON/OFF switch (green = ON, amber = OFF).

4 - CABIN MONITOR ON/OFF− Pressing the “CABIN MONITOR ON/OFF” button can turn ON

or OFF the forward 15” LCD monitor (green = ON, amber =OFF).

5 - CABIN VIDEO SELECTION− The “CABIN VIDEO SELECT” button can be used to select

“VCR”, “DVD1” or “DVD2”, which will be displayed on the LEDdisplay. The VCR/DVD video will appear clear and sharp on theindividual 6.5” monitor installed at the seat location.

6 - SPEAKER AUDIO SELECTION− Using the “SPEAKER AUDIO SELECT” switch, change the

audio until the VCR, DVD1 or DVD2 audio can be heard fromthe forward cabin speakers. The audio source selected will bedisplayed on the LED display.

7 - SPEAKER ON/OFF− Pressing the “SPEAKER ON/OFF” button can turn the forward

cabin speakers ON or OFF (green = ON, amber = OFF).


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9 - PERSONAL VIDEO SELECTION− The “PERSONAL VIDEO SELECT” button can be used to select


10 - CABIN TEMPERATURE CONTROLLER− Pressing the “CABIN TEMP CONTROL” button presents the

approximate cabin temperature (°F). With the “UP” and“DOWN” buttons cabin temperature can be adjusted.

11 - TABLE LIGHT− The table light can be turned ON or OFF by pressing the

“TABLE LIGHT” ON/OFF switch (green = ON, amber = OFF).

12 - AFT DOWNWASH LIGHTS− The aft cabin downwash lights can be turned ON or OFF by

pressing the “AFT DNWASH LIGHTS” ON/OFF switch (green =ON, amber = OFF).

13 - FORWARD DOWNWASH LIGHTS− The forward cabin downwash lights can be turned ON or OFF by

pressing the “FWD DNWASH LIGHTS” ON/OFF switch (green= ON, amber = OFF).

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FORWARD VIP SEAT PASSENGER CONTROL PANEL


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AFT VIP CONTROL PANEL (OPTIONAL)

The aft VIP control panel is located in the left-hand side panel near theaft VIP seat (optional). There is a headset jack andvolume/temperature “UP” and “DOWN” control buttons. The audio willcorrespond to the chosen video selection, however, it can be selectedseparately from the video. This panel has the ability to control thefollowing items:

1 - FORWARD UPWASH LIGHTS− The forward cabin upwash lights can be turned ON or OFF by

pressing the “FWD UPWASH LIGHTS” ON/OFF switch (green =ON, amber = OFF).

2 - AFT UPWASH LIGHTS− The aft cabin upwash lights can be turned ON or OFF by

pressing the “AFT UPWASH LIGHTS” ON/OFF switch (green =ON, amber = OFF).

3 - READING LIGHT− The reading light can be turned ON or OFF by pressing the

“READ LIGHT” ON/OFF switch (green = ON, amber = OFF).

4 - SPEAKER AUDIO SELECTION− Using the “SPEAKER AUDIO SELECT” switch, change the

audio until the VCR, DVD1 or DVD2 audio can be heard fromthe forward cabin speakers. The audio source selected will bedisplayed on the LED display.

5 - SPEAKER ON/OFF− Pressing the “SPEAKER ON/OFF” button can turn the forward

cabin speakers ON or OFF (green = ON, amber = OFF).



7 - PERSONAL VIDEO SELECTION− The “PERSONAL VIDEO SELECT” button can be used to select


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8 - CABIN TEMPERATURE CONTROLLER− Pressing the “CABIN TEMP CONTROL” button presents the

approximate cabin temperature (°F). With the “UP” and“DOWN” buttons cabin temperature can be adjusted.

9 - TABLE LIGHT− The table light can be turned ON or OFF by pressing the

“TABLE LIGHT” ON/OFF switch (green = ON, amber = OFF).

10 - AFT DOWNWASH LIGHTS− The aft cabin downwash lights can be turned ON or OFF by

pressing the “AFT DNWASH LIGHTS” ON/OFF switch (green =ON, amber = OFF).

11 - FORWARD DOWNWASH LIGHTS− The forward cabin downwash lights can be turned ON or OFF by

pressing the “FWD DNWASH LIGHTS” ON/OFF switch (green= ON, amber = OFF).


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AFT VIP SEAT PASSENGER CONTROL PANEL

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ESCUTCHEONSThe escutcheons are provided to support the gasper air outlets, theoxygen box assemblies, the speakers and the reading lights.

There are two types of escutcheons. The first assembly comprises tworeading lights, a speaker, and a speaker grill, while the second onecomprises an oxygen box assembly installed in the middle of theescutcheon, a reading light, and a gasper air outlet. In the left-handand right-hand escutcheons assemblies, the reading light and thegasper air outlet are positioned to establish symmetry between bothsides of the aircraft.

The escutcheons are installed in the left-hand and right-hand valancepanels in sets of three units each, one first escutcheon type betweentwo second escutcheon type, above the seats of conference table andexecutive tables, the executive divan seat, and the credenza, forpassengers comfort.

A second escutcheon type is installed in the lavatory valance panel.

The escutcheons provide the following services:

1 - READING LIGHTS.

2 - AIR GASPER.

3 - OXYGEN MASKS DISPENSING.

4 - LOUDSPEAKER FOR INTERNAL COMMUNICATION.


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1 - READING LIGHTS

− Press the “READ LIGHT” button on the passenger seat controlpanel to turn the reading lights on. To control light direction,push the reading light in the direction illumination is desired.

2 - AIR GASPER

− Turn the nozzle that protrudes from the ball assembly to controlthe airflow volume. To control airflow direction, push the nozzlein the direction in which the airflow is desired.

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ESCUTCHEONS - COMPONENT LOCATIONS


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CLOSETSThe LH (left-hand) forward closet, LH and RH (right-hand) closets,entertainment cabinet and credenza are used for stowage ofpassenger garments, miscellaneous items, life rafts, pillows, blankets,magazines, plug-in monitors and as a storage area for emergencyequipment, umbrella, entertainment equipment and 110 V AC outlets.

The LH forward closet, LH and RH closets, entertainment cabinet andcredenza are made of lightweight honeycomb panels and the exterioris covered with a decorative finish and laminate. The LH forward closetis located in the forward section of the passenger cabin area. The LHand RH closets are located in the forward section, next to the pocketdoor. The entertainment cabinet is located in the forward section infront of the galley. The credenza is located in the mid-section of thepassenger cabin area.


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COMPONENTS

LH FORWARD CLOSET

The LH forward closet is composed of two hinged door assemblies,coat rod, life raft storage, and umbrella storage. The LH forward closetcan storage a total of six umbrellas shelves. The interior of LH forwardcloset is accessed by means of two hinged door assemblies. Thedoors are held closed with latches.

The 15 inch LCD monitor and the MHR infrared control system may beinstalled at the LH forward closet depending on the interiorconfiguration.

The figure on the next page presents the LH forward closetcomponents location.

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LH FORWARD CLOSET - COMPONENT LOCATIONS


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LH AND RH CLOSETS

The LH and RH closets have a hinged door assembly and a coat rod.The door is held closed with a latch. These closets offer room for coatsand miscellaneous storage. The LH closet contains a 15 inch LCDmonitor, and the MHR infrared control system.

The RH closet is comprised of optional drawers, floor level warmoutlets and ECS/decompression airflow outlets.

The figure on the next page presents the LH and RH closetcomponents location.

NOTE: The LH and RH closets are only available for some passengercabin layouts.

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LH AND RH CLOSETS - COMPONENT LOCATIONS


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CREDENZA

The credenza has a magazine rack assembly and according to theinterior arrangement there are three or four hinged door assembliesand eight or ten drawer assemblies installed. The doors and drawersare held closed with latches.

The credenza includes escape path light, room for soft storage (pillowsand blankets), miscellaneous items, extra china and flatware storage,printer/fax machine, magazine, headphone, and life raft storage. It alsopossesses space provisions for a 10.4 inch pop-up LCD monitor, 110V AC outlet for printer/fax machine, floor level warm air outlets andECS/decompression airflow outlets.

The figures on the following pages present the credenza´s componentslocation.

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CREDENZA - COMPONENT LOCATIONS (WITH THREE HINGEDDOORS AND EIGHT DRAWER ASSEMBLIES)


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01 JUNE 20, 2002

CREDENZA - COMPONENT LOCATIONS (WITH FOUR HINGEDDOORS AND TEN DRAWER ASSEMBLIES)

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ENTERTAINMENT CABINET


The entertainment equipment rack, installed in the forward section ofthe passenger cabin area, has two storage compartments foremergency equipment, a vented storage compartment for AC-powerstatic inverters and laptop controller interface, entertainment circuitbreakers panel compartment, systems compartments, two DVDplayers, a video player, Airshow Genesys, SATCOM, clear ice, anentrance control panel, entertainment control panel, a media storagerack, protective breathing equipment (PBE), aisle lights and an area ofECS/decompression air outflow.

Three hinged-door assemblies and a removable panel give access toits interior. The doors are held closed with latches.

The following list presents the main components on the entertainmentcabinet, and the figure provided on the next page shows the location ofthese components.

1 - MEDIA STORAGE RACK2 - MASTER ENTERTAINMENT CONTROL PANEL3 - ENTRANCE CONTROL PANEL4 - MULTI-REGION DIGITAL VIDEO DISK PLAYER5 - HANDSET6 - MULTI-STANDARD VIDEO CASSETTE RECORDER7 - FIRE EXTINGUISHER8 - CLEAR ICE PROCESSOR9 - AUDIO AMPLIFIER10 - TEMPERATURE CONTROLLER11 - MH PROGRAMMMING INTERFACE12 - PAX AC STATIC INVERTER13 - MH INTERFACE14 - PBE/SMOKE HOOK15 - PORTABLE OXYGEN CYLINDER16 - CRASH AXE WITH LEATHER POUCH17 - FIRST AID KIT18 - FLASHLIGHT19 - ENTERTAINMENT CABINET CIRCUIT BREAKER PANEL20 - MH CABIN CONTROLLER21 - SATCOM22 - AIRSHOW DIGITAL INTERFACE UNIT23 - MH ENTERTAINMENT CONTROL PANEL

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ENTERTAINMENT CABINET - COMPONENT LOCATIONS(AIRPLANES UP TO S/N 685 (INCLUSIVE))


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FOR AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENTAIRFRAMES.

The entertainment equipment rack has a storage compartment foremergency equipment, a vented storage compartment for AC staticpower inverter, a laptop controller interface, an entertainment circuitbreakers panel compartment, system compartments, two DVD players,video player, airshow genesys, SATCOM, an entrance control panel,an entertainment control panel, a media storage rack, a PBE, and acurtain storage compartment. Three hinged-door assemblies and aremovable panel give access to its interior. The doors are held closedwith latches. The entertainment equipment rack is installed in theforward section of the passenger cabin area.

The following list presents the main components on the entertainmentcabinet, and the figure provided on the next page shows the location ofthese components.

1 - MASTER ENTERTAINMENT CONTROL PANEL2 - CD STORAGE RACK3 - VIDEO STORAGE RACK4 - ENTRANCE CONTROL PANEL5 - FIRE EXTINGUISHER6 - PROTECTIVE BREATHING EQUIPMENT7 - PORTABLE OXYGEN CYLINDER AND MASK ASSY8 - MEDICAL KIT9 - DIGITAL VIDEO DISK PLAYER10 - VIDEO CASSETE PLAYER11 - CIRCUIT BREAKERS PANEL12 - AUDIO AMPLIFIER 113 - COOLING FAN14 - MH ENTERTAINMENT CONTROLLER15 - SATELLITE DATA UNIT16 - AIRSHOW DIGITAL INTERFACE UNIT17 - AUDIO AMPLIFIER 218 - MH CABIN CONTROLLER19 - MHP LAPTOP CONTROLLER20 - MH INTERFACE CABIN MANAGEMENT21 - MHR IRS22 - TEMPERATURE CONTROLLER23 - AC STATIC POWER INVERTER

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ENTERTAINMENT CABINET - COMPONENT LOCATIONS(AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENT

AIRFRAMES)


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ENTRANCE CONTROL PANEL(AIRPLANES UP TO S/N 685 (INCLUSIVE))

The entrance control panel controls the main door and entry courtesylights.

The ENTRY LIGHTS ON/OFF push-button switch can turn ON or OFFthe courtesy and stair lights. Pressing the ENTRY LIGHTS AUTOpush-button switch, the courtesy and the stair lights will be ON whenthe main door is open.

The DOOR CLOSED push-button switch commands the main door toclose and the DOOR BLOCKED indication shows when the main dooris hydraulically blocked.

1 - ENTRY LIGHTS ON/OFF

− Press ENTRY LIGHTS ON/OFF to turn the courtesy and thestair lights ON/OFF.

2 - ENTRY LIGHTS AUTO

− Press ENTRY LIGHTS AUTO to turn the courtesy and the stairlights ON when the door is open.

3 - DOOR CLOSED

− Press DOOR CLOSED button to command the main door toclose.

4 - DOOR BLOCKED

− When the DOOR BLOCKED indication becomes amber themain door actuator hydraulic line remains pressurized after doorclosing. In this case the main door is hydraulically blocked.Otherwise, when it becomes green the hydraulic actuator line isdepressurized.

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ENTRANCE CONTROL PANEL - CONTROLS AND INDICATORS


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ENTRANCE CONTROL PANEL(AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENTAIRFRAMES)

The entrance control panel controls the main door and entry lights.Theairstair lights on/off push-button switch controls the stair lights; the steplight on/off push-button switch controls the step light; and the area lighton/off push-button switch controls the door area light. The airstair doorclose push-button switch controls the main door to close and the doorblocked push-button switch commands the main door to block.

1 - AIRSTAIR DOOR CLOSE

− Controls the main door to close.

2 - DOOR BLOCKED

− Controls the main door to block.

3 - AIRSTAIR LIGHTS ON/OFF

− Controls the stair lights on/off.

4 - STEP LIGHT ON/OFF

− Controls the step light on/off.

5 - AREA LIGHT ON/OFF

− Controls the door area light on/off.

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MASTER ENTERTAINMENT CONTROL PANEL

6 - DISPLAY

− Monitors source selection

7 - SPEAKER SELECTION INDICATION

− Indicates if the speaker is selected.

8 - FORWARD VIDEO SELECTION INDICATION

− Indicates if the forward video is selected.

9 - AFT VIDEO SELECTION INDICATION

− Indicates if the aft video is selected.

10 - TEMPERATURE

− Displayed in degrees Fahrenheit.

11 - MAP DISPLAY MODE; (MAP, AUTO, LOGO, INFO)

− Selects the map modes of the Airshow.

12 - AFT LCD ON/OFF SWITCH

− When alternately pressed, turns ON or OFF the aft LCD.

13 - FORWARD LCD ON/OFF SWITCH

− When alternately pressed, turns ON or OFF the forward LCD.

14 - SPEAKER ON/OFF SWITCH

− When alternately pressed, turns ON or OFF the speaker.

15 - VOLUME UP KEY

− Increases the volume on the audio speakers.

16 - VOLUME DOWN KEY

− Decreases the volume on the audio speakers.


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ENTRANCE/MASTER ENTERTAINMENT CONTROL PANEL(AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENT

AIRFRAMES)


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PARTITIONSThe partitions are used to separate the cockpit from the passengercabin area, the forward galley area from forward passengercompartment, the forward passenger compartment from the aftpassenger compartment, the lavatory from the passenger cabin area,and the lavatory from the baggage compartment.

The EMB-135BJ is equipped with the following partitions:Cockpit/passenger cabin partition, which separates the cockpit fromthe passenger cabin area with controllable access door between thetwo areas; the pocket door partition, which separates the forwardgalley area from forward passenger compartment and contains asliding pocket door; the cabin partitions, which separates the forwardpassenger compartment from the aft passenger compartment; thetoilet partition, which separates the lavatory from the passenger cabinarea and incorporates a swing door; and the lavatory/baggagecompartment partition, which separates the lavatory from the baggagecompartment, and is also provided with a swing door that permitspassage from one area to another.

The following figure shows the location of the partitions.


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PARTITIONS LOCATION


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COMPONENTS

COCKPIT/PASSENGER CABIN PARTITION

The cockpit/passenger cabin partition possesses a cockpit door, LHand RH partitions.

The cockpit door assembly is composed of a blowout panel assemblyand a door assembly. The door assembly contains a locking latchassembly and a viewer. Two strap assemblies keep the blowout panelassembly attached to the door in the event of a significant pressureincrease or decrease in the aircraft. The peephole allows the cockpitoccupants to see through the door assembly. The locking latchassembly engages a striker on the LH partition to secure the doorwhen it is closed.

The cockpit/passenger cabin LH partition contains a striker and anelectrical installation. The electrical installation possesses a harnessassembly.

The cockpit/passenger cabin RH partition is provided with an accesspanel assembly that gives access to the relay box.

One blue light can be provided above the cockpit door to advise whenthe cockpit is to be shut out. This light is commanded through theSTERILE light switch located at the overhead panel.

The figure on the next page presents the location of thecockpit/passenger cabin partition components.

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COCKPIT/PASSENGER CABIN PARTITIONCOMPONENT LOCATIONS


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POCKET DOOR PARTITION

The pocket door partition is a decorative lightweight panel assemblyinstalled with fasteners to the forward galley.

The function of the pocket door is to reduct noise level and to provideprivacy to the passenger cabin area.

There are two types of pocket door partitions, which are the single anddual sliding pocket door.

Both pocket doors open by a lateral sliding movement and have alocking mechanism to lock the door in the open position for takeoffsand landings.

The figures on the next pages present the two types of pocket doorsand their partition components.

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SINGLE POCKET DOOR PARTITION - COMPONENT LOCATIONS


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DUAL POCKET DOOR PARTITION - COMPONENT LOCATIONS


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CABIN PARTITION

The cabin partition has a LH partition, a RH upper partition, and a RHlower partition made of decorative honeycomb panels. They are alsolined with a gloss and veener coat that gives this partition a decorativefinish. The panels are tapered at the top to provide an “open”appearance. The LH partition is provided with “EMERGENCY EXIT”,“NO SMOKING”, and “FASTEN SEATS BELTS” signs on both sides.

The RH partition is removable.

The figure on the next page presents the location of cabin partitioncomponents.

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CABIN PARTITION - COMPONENT LOCATIONS


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TOILET PARTITION

The toilet partition has a LH partition, a RH partition, and a swing doormade of decorative honeycomb panels installed at the rear of thepassenger cabin area. The toilet partition is lined with a gloss andveener coat. The doorknobs and latch are gold plated. The doubleswing door is a two-way operable door and contains a door pop-upheader, and a doorknob with “VACANT/OCCUPIED” sign controlled bya slide bolt that can be overridden from the cabin side.

The LH partition includes “NO SMOKING” and “FASTEN SEATBELTS” signs.

The figure on the next page presents the location of the toilet partitioncomponents.

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TOILET PARTITION - COMPONENT LOCATIONS


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LAVATORY/BAGGAGE COMPARTMENT PARTITION

The lavatory/baggage compartment partition is composed of LH andRH partitions. The LH partition contains a baggage compartment doorand a baggage lights control panel that is turned ON/OFF in thelavatory.

The baggage lights control panel controls the two incandescent lightsof the baggage compartment.

The forward surface, including the door, lined with a gloss and veneercoat that gives this partition a decorative finish. The aft side of thelavatory/baggage compartment partition is lined with gross-point fabric.The baggage compartment door contains a latch mechanism, ablowout panel assembly, a guard assembly, and a peephole (one-wayinto the baggage area). The latch mechanism allows the door to beopened from the baggage compartment interior.

The guard assembly is made of a metallic structure and is designed tocatch the blowout panel assembly should this panel be blown-out.Within the guard assembly, there are two lanyards that keep theblowout panel attached to the door if a remarkable pressure differenceoccurs between the lavatory and the baggage compartment.

The lavatory/baggage compartment partition also contains amicroswitch that provides an EICAS message indicating door openstatus.

The figure on the next page presents the location of thelavatory/baggage compartment partition components.

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LAVATORY/BAGGAGE COMPARTMENT PARTITIONCOMPONENT LOCATIONS


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1 - RELAY BOX ACCESS PANEL

− To operate the relay box panel on the LH partition located at thecockpit/passenger cabin partition, access through the life vestcompartment is required.

− To operate the relay box panel on the RH partition located at thecockpit/passenger cabin partition, access through the forwardgalley is required.

2 - POCKET DOOR

− To open the pocket door, release the latch and smoothly slide itthrough the tracks without binding or hesitation to the RH side.

− To close the pocket door, smoothly slide it through the trackswithout binding or hesitation to the LH side and fit the latch.

The figure on the next page presents the operation of the pocket door.

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POCKET DOOR - OPERATION


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WATER AND WASTEThe water and waste system has three sub-systems: the water supplysystem, the water/waste drain system, and the toilet waste system.

In the lavatory compartment there is a washbasin unit and a toiletassembly, while in the forward galley there is a washbasin unitinstalled. Some airplanes may be equipped with a forward lavatory,where there is a washbasin unit and a toilet assembly that are suppliedby an independent water system.

Each washbasin unit furnishes drinking water through pressurizedwater lines, and drains wastewater overboard by differential pressure.

The water/waste drain system routes wastewater from the lavatory andgalley washbasin units and other sources to the overboard water/wastedrain ports.In the case of forward lavatory, the water/waste is drained to theforward lavatory waste tank. This waste tank is serviced by inside theforward lavatory.

The waste tank unit contains a liquid disinfectant that cleans the bowlthrough cycling. The system also has an arrangement for odorexhausting from the toilet assembly.

The toilet waste system provides sanitary means to collect toilet wastefor proper disposal when the aircraft is on the ground.

WATER

FOR AIRPLANES UP TO S/N 685 (INCLUSIVE):

The water system stores water for drinking and washing purposes in apressurized and freeze-protected tank. It assures no contamination ofthe water by being made up of stainless-steel components and usingtwo water filter units installed near the water tank.

It supplies the washbasin with water at ambient cabin temperature andheated water through the supply lines, which connect the tank to thewashbasin faucets, every time and while the faucets are in use.

The water system control panel is installed in the forward galley toprovide status indication and control for the water system.

Water servicing is done through the external water service panel, onthe lower rear right side of the wing-to-fuselage fairing, which allowsthe supply of water to the tank and drain, if necessary with heatednipples with caps (fill and overflow), switches, drain valve, and acontrol cable.


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The nipples mounted on the external service panel are directlyconnected to the tank through stainless-steel tube assemblies andhoses.

There is a back drain valve that provides means to drain the watersystem during ground servicing. A drain switch located on the waterservice panel actuates the valve.

Each washbasin, one in the galley and the other in the lavatorycompartment, has a manual isolation shutoff valve, a water heater, anda hot/cold water faucet. In the event of a water leak, the washbasinscan be shut-off by closing the isolation valves. The water heatersprovide hot water for the faucets.

In order to prevent freezing during high-altitude, long-duration flights,there is an external electric heater blanket on the water storage tankand two types of electric heaters for the water distribution lines, whichare the inner-line heater and the heated hose.

The inner-line electric heater requires 115 V AC - 400 Hz electricalpower, which is supplied from a heater controller. The inner-line heateris inserted into the water distribution line through an interface fitting.

The heated hose is flexible and has an integrated external electricalheating element. The heated hose requires 28 V DC power and aremotely located thermostat switch.

Electric heaters are installed on the water tank drain valve. Theheaters require 28 V DC power and are controlled by the twothermostat switches that control the heated hoses.

Electric heaters are also installed on the fill and the vent-overflow portson the water service panel.

FOR AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENTAIRFRAMES:

Before any service is performed on the water system, the operatorneeds to ensure that the system is de-energized at the water controlpanel located in the upper aft Galley, upper right hand cabinet door.

NOTE: It is recommended that the Water System be drained if theaircraft is expected to stay overnight in an environment belowthe freezing point of water.

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The water system stores water for drinking and washing purposes intwo pressurized and freeze-protected tanks, one located in the upperforward galley (with 15 liters - 3.9 US gallons of capacity), the other inthe upper portion of the aft lavatory (with 35 liters - 9.2 US gallons ofcapacity). Both tanks employ stainless-steel components and waterfilter units, which inhibit water contamination.

The galley and aft lavatory tanks supply the washbasin with water atthe ambient cabin temperature and heated water through the supplylines, which connect the tank to the washbasin faucets, every time andwhile the faucets are in use.

The water system control panel is installed in the forward galley toprovide status indication and control for the water system.

Water servicing is done through two external water service panels, oneon the lower rear right side of the wing-to-fuselage fairing and the otheron the lower front right side of the wing. Both allow the supply of waterto the tanks and drain, if necessary with heated nipples with caps (filland overflow), switches, drain valve, and a control cable.For the optional forward lavatory, the Water System is serviced frominside the forward lavatory compartment. Water is stored in a watertank (with 10 liters - 2.6 US gallons of capacity) that supplies theforward lavatory washbasin.

The nipples mounted on each external service panel are directlyconnected to the associated tank through stainless-steel tubeassemblies and hoses.

There is a back drain valve that provides the means to drain the watersystem during ground servicing. The valve is actuated by a drainswitch located on the water service panel.

Each washbasin, one in the galley and the other in the lavatorycompartment, has a manual isolation shutoff valve, a water heater, anda hot/cold water faucet. In the event of a water leak, the washbasinscan be shut-off by closing the isolation valves. The water heatersprovide hot water for the faucets.

In order to prevent freezing during high-altitude, long-duration flights,there is an external electric heater blanket on each water storage tankand two types of electric heaters for the water distribution lines, whichare the inner-line heater and the heated hose.


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The inner-line electric heater requires 28 V DC electrical power, whichis supplied from a heater controller. The inner-line heater is insertedinto the water distribution line through an interface fitting.

The heated hose is flexible and has an integrated external electricalheating element. The heated hose requires 28 V DC power and aremotely located thermostat switch.

Electric heaters are installed on each water tank drain valve. Theheaters require 28 V DC power and are controlled by the thermostatswitches that control the heated hoses.

Electric heaters are also installed on the fill and the vent-overflow portson each water service panel.

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WATER SYSTEM - CONTROL AND SERVICE PANELS


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WASTE DISPOSAL

The waste system consists of an electrically-operated self-containedrecirculation toilet unit, which collects and stores human waste in aninternal holding tank. Adequate chemical products are used to disinfectand deodorize the waste holding tank.A vent line connecting the waste holding tank to the exterior performsits ventilation (odors exhaust) by means of differential pressure.Toilet flushing is initiated by pressing and releasing the flush buttonadjacent to the toilet. This button actuates a motor-driven pump andfilter, which delivers flushing fluid for a pre-timed interval.A restrictor at the bowl bottom prevents waste material return when it iscarried directly to the tank.A waste service panel on the lower rear right side of the fuselage isequipped with a control cable, a waste drain valve and a rinse nipplewith cap, and allows the waste system to be serviced.

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WASTE DISPOSAL - WASTE SERVICE PANEL LOCATIONAND SCHEMATIC DIAGRAM


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The optional forward lavatory has its own separated waste system,which consists of an electrically-operated self-contained recirculationtoilet unit, which collects and stores human waste in an internal holdingtank. Adequate chemical products are used to disinfect and deodorizethe waste holding tank.

The forward lavatory waste servicing is done by taking out the wastetank from inside the lavatory.

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FORWARD LAVATORY - WASTE DISPOSAL


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AIRSTAIR MAIN DOOR

The aircraft is provided with one main entry door located at the leftforward fuselage section.The main door, incorporating folding airstairs, is hinged at its loweredge. The door is raised under normal operation conditions by twohydraulic door actuators powered by hydraulic system 1 or by anaccumulator with sufficient capacity for three complete door operationcycles.The door opening operation is manual. The hydraulic circuit dampingfunction allows a smooth operation when the door is lowered.The system may be controlled from inside or outside the airplane,respectively through the internal main door control panel, which is partof the entrance control panel, or through the exterior main door controlpanel.The internal main door control push-button DOOR CLOSED is locatedin the internal main door control panel. The external main door controlpush-button PRESS CLOSE DOOR is located behind an access panelon the lower left-hand side of the door´s lower edge.The door may also be closed and locked by raising it manually, by anoutside ground attendant, actuating either the inner or the outerhandle.An alternative main door opening valve is provided in the cockpit toallow the main door to be lowered if it is blocked by hydraulic systempressure (selector valve failure).

NOTE: No more than three persons should simultaneously bestanding on the doorsteps.

EICAS MESSAGE


WARNING MAIN DOOR OPNMain door is open or not properlylocked either on the ground withengine 1 running or in flight.


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1 - EXTERIOR MAIN DOOR CONTROL BUTTON− When pressed, a solenoid valve is energized, allowing hydraulic

power to raise the main door.

2 - INTERIOR MAIN DOOR CONTROL BUTTON - DOOR CLOSED− When pressed, a solenoid valve is energized, allowing hydraulic

power to raise the main door.− The DOOR BLOCKED indication becomes amber on the

internal main door control panel when the main door actuatorhydraulic line remains pressurized after door closing. In thiscase the main door is hydraulically blocked.

3 - ALTERNATE MAIN DOOR OPENING VALVE (EMERGENCYDISCHARGE VALVE)− Open the access panel of the emergency discharge valve.− Turn the emergency discharge valve button clockwise to

decrease the high pressure in the pressure/return line of theactuators.

− When actuated for 2 minutes, it depressurizes the door closeline, allowing the main door to be lowered when blocked byhydraulic system pressure, provided Hydraulic System 1 isdepressurized.

− The DOOR BLOCKED indication becomes green on the internalmain door control panel, and the main door opens smoothly.

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AIRSTAIR MAIN DOOR CONTROLS AND INDICATORS


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AIRSTAIR DOOR OPERATION (INSIDE CABIN)

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AIRSTAIR DOOR OPERATION (OUTSIDE CABIN)


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MAIN DOOR ACOUSTIC CURTAIN

The airplane is equipped with an acoustic curtain at the main doorarea. The acoustic curtain reduces noise level in the forwardpassenger cabin area when it is installed.

NOTE: - The acoustic curtain must be stowed for takeoff andlanding.

- The acoustic curtain should be installed during flights forpassenger comfort.

- The acoustic curtain should be rolled-up with the ultra-leather facing outward. Thus, in case of rain, snow, wind orother weather conditions, the ultra-leather will be theexposed material.

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MAIN DOOR ACOUSTIC CURTAIN


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ACCESS DOORS AND HATCHESThere are some doors that provide access to several airplane systemsin the pressurized and in the non-pressurized compartments.

The lavatory/baggage compartment door provides ground or in-flightaccess to the baggage compartment from the lavatory.

The control rigging door provides access to the airplane´s systemslocated at the pressurized compartment below the cockpit floor. Therear electronic compartment door provides access to the electroniccomponents located in this pressurized compartment.

The forward electronic compartment door, the pressure-fueling controlpanel door and the landing gear doors provide access to the non-pressurized compartments.

The baggage door, on the aft left-hand side of the fuselage providesexternal access to the baggage compartment.

The aircraft is provided with three emergency exits for crew andpassenger emergency evacuation, which are as follows 1) Two cockpitemergency exits, and 2) One passenger cabin emergency exit, locatedover the wing on the right side of the fuselage. This emergency exit isone of the two passenger cabin emergency escape hatches, locatedover the wings. The other escape hatch, located on the left side of thepassenger cabin, is permanently locked and cannot be used as anemergency exit. This hatch, however, can be removed formaintenance purposes. The pilot and copilot direct-vision windows areused as the cockpit emergency exits by unlatching and removing theminward the cockpit.

All doors have a warning system, except for the landing gear doors andthe cockpit emergency exits.

Access doors and hatches locations are presented in the figures on thefollowing pages.


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ACCESS DOORS AND HATCHES LOCATIONS

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ACCESS DOORS AND HATCHES LOCATIONS 2


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BAGGAGE DOOR

The baggage door is located on the rear left side of the fuselage,below the pylon, and is manually operated from the outside. It isprovided with a locking mechanism controlled by an external handlethat is stowed in a recess in the mid-lower portion of the door. Thedoor is provided with a depressurization vent that allows openingoperations.

EICAS MESSAGE


CAUTION BAGGAGE DOOR OPNBaggage door open or notproperly locked.

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BAGGAGE DOOR OPERATION


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COMPARTMENT HATCHES

A number of access doors and hatches for different aircraft systemscan be found along the fuselage.

The compartment hatches provide access for servicing the airplane´ssystems and equipment.

The ventral cockpit access hatch is located under the fuselage,providing access to the fuselage pressurized compartment.

The forward electronic compartment access hatch is inside the noselanding gear wheel well.

The rear electronic compartment access hatch is located on the rearright side of the fuselage. This hatch provides access to the airplane´spressurized area, which contains the rear electronic compartment,rudder autopilot servo, rudder control cables and electrical harness,stabilizer electrical harness and elevators control cables.

An unlocked condition of any compartment hatch causes a singlecaution message on EICAS. In addition, the MFD indicates an openhatch (es) condition through a graphical representation.

EICAS MESSAGE


CAUTION ACCESS DOORS OPNAt least one compartmentaccess hatch is open or notproperly locked.

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REFUELING PANEL ACCESS DOOR

The refueling panel access door is located on the forward right side ofthe wing-to-fuselage fairing (refer to Section 2-8 – Fuel System).

Opening the refueling panel access door causes an EICAS cautionmessage. In addition, the MFD indicates an open-door conditionthrough a graphical representation.

EICAS MESSAGE


CAUTION FUELING DOOR OPNRefueling panel access dooropen or not properly closed.

LAVATORY/BAGGAGE COMPARTMENT ACCESSDOOR

The lavatory/baggage compartment access door is located in thelavatory/baggage compartment partition, and provides in-flight accessto the baggage compartment.

This access door must be closed for takeoffs and landings.

Opening of the lavatory/baggage compartment access door causes anEICAS caution message. In addition, the MFD indicates the open-doorcondition through a graphical representation.

EICAS MESSAGE


CAUTION BAGG ACCESS OPNLavatory/Baggagecompartment access dooropen or not properly closed.


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EMERGENCY EXIT HATCHES

One passenger cabin emergency escape hatch is located over the left-hand wing. (Refer to Section 1-10 – Emergency Information).

DOORS AND HATCHES INDICATION ON MFD

The DOORS section of the Takeoff System Page of the MFD consistsof a graphical representation of the airplane (white) with squareslocated along the fuselage to denote the various doors and hatchesthat are monitored.

If a door or hatch is ajar, the associated graphical square will changefrom green to red and a red DOOR OPEN inscription will be presented,boxed in red, in the lower left corner of the DOORS section.

The following doors and hatches are monitored for status:

− Main door;

− Baggage door;

− Lavatory/Baggage compartment access door;

− Fueling panel access door;

− Rear electronic compartment access hatch;

− Forward electronic compartment access hatch;

− Under cockpit access hatch;

− Emergency exit hatch.

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DOORS AND HATCHES INDICATION ON MFD


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FORWARD LAVATORY DOORS

INTRODUCTION

The forward lavatory compartment is located in the service area, whichis mainly used by the crewmembers. The forward lavatory offersadequate conditions for the crew´s personal hygiene and amenities.

FORWARD DUAL POCKET DOOR

The forward dual pocket door partition is divided into the LH forwardpocket door partition and the RH forward pocket door partition. Eachside comprises a door hinge fitting, door latch, and an indicator latch.The forward dual pocket door partition structurally separates theforward lavatory/service area and the seating area of the passengercompartment.

In case of the dual pocket door becomes jammed, push or kick at thecenter until door breaks free. To ensure a free pathway restow thedoor as follows:− Swing back, slowly, each door leaf perpendicular to the aisle.− Lift up each leaf until it locks.− Stow each leaf in the pocket as in normal operation.

BI-FOLD FOWARD LAVATORY DOOR

The door on the fwd lavatory is a bi-fold door that when opened 90°locks onto the galley, across the aisle, closing off the forward part ofthe service area. The bi-fold door is provided with means to unlock thedoor from outside.

When the bi-fold lavatory door separetes at the center, in order toensure a free pathway, swing both RH and LH leaves back and securethem as follows:

RH door:− Locate lanyard underneth galley toe kick− Attach the lanyard hook to the bracket located underneath the RH

door leaf.

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LH door:

1 - Engage door arm on T-rail.2 - Remove bolt.3 - Rotate locking arm and align with hole A.4 - Install bolt and nut.

OPERATION

To use the fwd lavatory, the forward dual pocket door and the bi-foldforward lavatory door must be operated as follows.

To enter:

− In the service area, slide the RH and LH side of the forward dualpocket door to close and latch them.

− The doors must close smoothly without binding or hesitation andthe latch will secure the door in the closed position.

− Unlatch and open the bi-fold forward lavatory door by sliding it tothe forward side until it reaches its path limit and release it fromits upper track.

− Unfold the door toward the galley and lock it to the galley fittingsusing the pin locks installed in the door´s upper and lower parts.

− Open the forward and aft close up panels using the knobs of thebi-fold door to close the gap between the upper part of the bi-fold door and the headliner.


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FORWARD LAVATORY BI-FOLD DOOR - OPERATION

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To exit:

− Close the forward and aft close up panels using the knobs of thebi-fold door.

− Unlock from the galley fittings the pin locks installed in the bi-fold door´s upper and lower parts and fold the bi-fold doortoward the forward lavatory until it fits the upper track.

− Slide the bi-fold door aft and lock it.− Slide the RH and LH sides of the forward dual pocket door to

open them.

The doors must open smoothly, without binding or hesitation, and thelatch will secure the door in the open position.

Cabin depressurization:

In the event of inadvertent cabin depressurization, the bi-fold door hastwo hinges installed with breakable pins that will segregate the door into two halves.

One half, fixed to the galley, will rotate 90° and the other half, fixed tothe lavatory structure, will also rotate to permit airflow between thepassenger cabin entrance area and the passenger cabin.

At the same time, the entire forward RH and LH pocket door panelsrotate, just one half panel being enough for the airflow to go through.

After the equalization of the pressure cabin , the bi-fold door must bereinstalled in the lavatory structure with two center spring rollers andthe door sliding fitted into position. This procedure permits keeping thebi-fold door stowed until it is repaired.

The pins will need to be replaced before any the lavatory operation.

In the case of accidental cabin depressurization, the lavatory hasunobstructed and redundant lower air paths that allow airflow from thepassenger cabin to the back of the lavatory (fuselage), therebyproviding quick cabin air pressure equalization. These air paths,located in the toe kick panel, are screened to avoid the ingestion andtrapping of unwanted objects falling into the area around thelavatory/galley.

In the aft lavatory, the aft pocket door provides the same operation asthat of the forward dual pocket door in case of depressurizationbetween areas in the aft part of the aircraft.


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AFT LAVATORYThe lavatory compartment is a modular unit that supplies the adequateconditions for the flight crew´s and passenger´s personal hygiene andamenities.

The lavatory is located in the aft section of the passenger cabin. Therear wall of the lavatory is covered by the passenger cabin/baggagecompartment partition.

The following figure presents the location of the lavatory´s maincomponents.


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LOCATION OF THE LAVATORY´S MAIN COMPONENTS


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VANITY ASSEMBLY

The vanity assembly is a bonded structure which houses the lavatorycloset assembly, the waste container and its automatic fireextinguisher, miscellaneous storage, vanity light shroud and passengersigns NO SMOKING/RETURN TO SEAT. Other accessories for thevanity assembly include a soap dispenser assembly, a mirror, a toiletpaper roll housing and potable water supply emergency shutoff valve,electric hot water heater, 110 V GFI outlet, vanity switch panel and abaggage lights control panel.

The baggage lights control panel is turned ON/OFF in the lavatory andis installed on the lavatory/baggage compartment partition.

The lavatory closet is composed of a storage compartment, coat rodand drawer.

TOILET SECTION COMPONENTS

The toilet components of the toilet assembly are the drip panassembly, the floor section assembly, the toilet floor lining bracketassembly, an electric flush toilet, a padded hinged toilet cover, thepotable water system shroud, lid assembly, a toilet tissue dispenser,and a valance panel.

The valance panel is composed of an escutcheon, in which there aretwo reading lights over the toilet assembly that can be turned ON/OFFby the reading lights switch on the lavatory control panel.

The figure on the next page presents the location of the vanityassembly and the toilet section components.

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LOCATION OF THE VANITY ASSEMBLY AND TOILET SECTIONCOMPONENTS


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LAVATORY ELECTRICAL INSTALLATION

There is a smoke detector installed on the ceiling section assembly.The upwash and downwash lights, two ballasts, an oxygen container,and one dimmer are on the valance.

The upwash and downwash lights turn to the lit condition when thelavatory door is closed, and they are controlled by a microswitch in thedoor.

An OCCUPIED, EXIT, and NO SMOKING/RETURN TO SEAT signsare installed on the forward face of the lavatory´s forward wall. The NOSMOKING/RETURN TO SEAT sign is operated from the cockpit. TheLAV CALL, reading lights and flush switches are installed on thelavatory control panel in the right-hand lavatory closet assembly. Whenthe LAV CALL switch is pressed, a tone will sound and the galleycontrol panel will illuminate the LAV indicator to call the flight attendant.

A flush switch assembly for toilet actuation is located next to the toiletpaper roll housing. When the toilet flush switch is pressed, the toiletflush cycle is initiated.

The reading light is installed on the valance panel over the toilet. It isilluminated with a LED module and the beam is adjustable. Pressingthe READ LIGHT labeled switch will turn the reading light ON or OFF.

The vanity assembly comprises the vanity light switch, the passengersign NO SMOKING/RETURN TO SEAT and the 110 V GFI outlet. Thevanity light is a fluorescent tube over the vanity mirror, which iscontrolled by the VANITY LIGHT ON/OFF labeled switch.

The figure on the next page presents the lavatory electrical installationcomponent locations.

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LAVATORY CONTROL PANEL

The lavatory control panel has the ability to control the following items:

1 - FLIGHT ATTENDANT CALL BUTTON− Press this button to send a signal to call the flight attendant.

2 - LAVATORY READING LIGHTS− When alternately pressed, turns the lavatory reading light ON or

OFF.

3 - TOILET FLUSH BUTTON− Press this button to flush the toilet after use.

VANITY CONTROL PANEL

The vanity control panel has the ability to control the following item:

1 - VANITY LIGHTS− When alternately pressed, turns the lavatory vanity lights ON or

OFF.

BAGGAGE LIGHTS CONTROL PANEL

1 - BAGGAGE COMPARTMENT LIGHTS− When alternately pressed, turns the baggage compartment

lights ON or OFF.

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LAVATORY, VANITY AND BAGGAGE COMPARTMENT LIGHTSCONTROL PANELS


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FORWARD LAVATORYThe forward lavatory is optional for airplanes from S/N 686 (inclusive)and subsequent airframes. The lavatory compartment is a modular unitthat supplies adequate conditions for the flight crew´s and passenger´spersonal hygiene and amenities.

This lavatory is located at the left forward section of the passengercabin.


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FORWARD LAVATORY


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TOILET SECTION COMPONENTS

The Forward Lavatory is installed in a cabinet that includes:− Vanity/toilet assembly with a bi-fold door.− LH forward-pocket door.− Forward lavatory electrical installation.

The rear wall of the lavatory is covered by the LH forward pocket doorcover partition and the front wall is attached to the entertainmentcabinet.The lavatory door is a bi-fold type. When swung open 90 degrees, itlocks into the galley, across the aisle, closing off the forward part of theservice area. The aft area of the lavatory is also closed when thelavatory is in use.The lavatory bi-fold door and the RH/LH forward pocket door functionas a blow-out panel in case of cabin decompression.The forward lavatory closet houses:

− Self-contained trash receptacle.− Lavatory light/accent light.− RTS/NS signs panel (RETURN TO SEAT/NO SMOKING).− Sink.− Faucet assembly.− Soap dispenser assembly.− Vanity mirror.− Toilet paper roll housing.− Potable water tank.− Water tank shut-off valve.− Electric hot water heater.− 110 V GFCI (GROUND FAULT CONTROL ISOLATION) outlet.− Lavatory indicator panel.− Miscellaneous storage compartment.− Self contained toilet.− Toilet seat.− Padded hinged toilet cover.− Lid assembly.

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VANITY ASSEMBLY AND TOILET SECTION -COMPONENT LOCATIONS


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LAVATORY ELECTRICAL INSTALLATION

The RTS/NS signs panel (RETURN TO SEAT/NO SMOKING) isinstalled on the vanity assembly and is operated from the cockpit.The flush switch is installed on the vanity assembly. When the toiletflush switch is pressed, the toilet flush cycle is initiated.The electrical installation also has a 110 V GFCI outlet installed on thevanity assembly.The lavatory light is installed at the top of the lavatory and an accentlight is installed on the vanity assembly. The lights are controlled by amicroswitch installed in the lavatory closet. Both lights are turned onwhen the lavatory bi-fold door is in the open position and they areturned off when the lavatory bi-fold door is in the closed position.


The lavatory has an indicator panel and a flush switch with thefollowing functions:

− RTS/NS signs panel (RETURN TO SEAT/NO SMOKING)These signs are installed on the vanity assembly and areoperated from the cockpit.

− FLUSH SWITCHThis push-button controls the toilet flush. When the toilet flushswitch is pressed, the toilet flush cycle is initiated.

− 110V OUTLETDetailed in the next figure.

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LAVATORY CONTROL PANEL - CONTROLS/INDICATORS


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PILOT AND PASSENGER CONVENIENCEITEMS

The Pilot and Passenger Convenience Items describe the systemsinstalled on the airplane to provide comfort, facilities and entertainmentfor passengers and crew members.

The systems described in this section are optional and some of themmay not be contained in your manual.

The Pilot and Passenger Convenience Items are:

- PC Power System.

- IFE - In-Flight Entertainment System.

- SATCOM - Satellite Communication System.

- Telephone System.


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PC POWER SYSTEM

The PC Power system is designed to supply passenger cabin ACelectric outlets with 115 V AC/60Hz power for Personal ElectronicDevices (PED), such as laptop computers.

A control switch in the In-flight Entertainment Panel (IFE), installed inthe Cockpit Overhead Panel, controls the single-phase AC staticinverter, located in the entertainment cabinet, which converts 28 V DCfrom Shed DC Bus 1 into 115 V AC/60 Hz. This control switch is thePC POWER push-button, which sends a ground signal to start the ACinverter operation.

The PC Power system starts to operate when the PC POWER push-button is set to the ON position (released). If the PC POWER push-button is set to the OFF position (pressed), the AC static inverteroutput is disabled and the striped bar in the push-button comes on.

If Shed DC Bus 1 is deenergized, the AC static inverter is alsodeenergized. Thus, the AC static inverter does not operate when theaircraft has only two or less generators providing power to the electricalsystem.

The output voltage and frequency are electronically controlled. The ACstatic inverter is provided with protection against input under/overvoltage, output under/over voltage, overcurrent, under/over frequency,input reverse polarity, and short-circuit.

The entertainment cabinet fuse box is installed behind theentertainment cabinet and is connected to the Shed DC Bus 1 througha fuse. An another fuse protects the AC static inverter generation.


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GROUND FAULT CIRCUIT INTERRUPTER

The Ground Fault Circuit Interrupter (GFCI) receptacle is a devicewhich provides personnel protection by interrupting a circuit where afault current to ground exceeds a predetermined value, and suppliespower to the AC electric outlets. It also permits localizing powerinterruption, provides convenient testing and resetting at the receptacleitself and minimizes nuisance tripping from extraneous causes. Thereare two GFCI installed in the aircraft that can be used as an outletitself. One is installed in the forward galley, which supplies power to theoutlets on the left side of the aircraft; and the other in the lavatory,which supplies power to the outlets on the right side of the aircraft.

NOTE: The use of the AC electric outlets is restricted during takeoffand landing. On ground, when the aircraft is energized withexternal power source (GPU), the AC static inverter output isavailable, and when the aircraft is energized with only two orless generators, the AC static inverter output is available only ifthe Shed Bus switch, on the electrical panel, is set to theOVRD position.

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IN-FLIGHT ENTERTAINMENT PANEL

The PC Power System is only activated or deactivated by releasing orpressing the PC Power Button on the IFE Panel.

1 - PC POWER BUTTON− Enables (released) or disables (pressed) the power supply to

the AC-outlet units.

− A striped bar illuminates in the button to indicate that it ispressed.

PC POWER SYSTEM BUTTON ON THE IFE PANEL


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IN-FLIGHT ENTERTAINMENT SYSTEM

The In-Flight Entertainment (IFE) system provides the passengers withreal-time flight information through the Airshow system, andvideo/audio entertainment through two DVD players, a VHS videocassette player and a CD changer.

The IFE control switch on the IFE Panel, installed in the CockpitOverhead Panel, controls the IFE system by activating the IFE masterrelay, which supplies power to the entertainment cabinet circuit breakerthrough Shed DC Bus 1.

The IFE system is supported by the Magic Happy (MH) Data Bus,which, aside the IFE, controls other airborne systems, such as thelighting system and the environmental control system. The MH digitalunits, such as the control units, are micro-processed andprogrammable by the MH Programming Interface that enables the useof a laptop with a proper configuration software.

The video generation on the IFE system is obtained through a videocassette player, two DVD players and the Airshow digital interface unit.The video signal is distributed to the eight 6.5” monitors and the 15”monitor via the MHE Entertainment Controller, which is commanded byunits such as the Master Entertainment control panel, the Forward VIPseat control panel, the Aft VIP seat control panel, the Passenger seatcontrol panels and a remote control unit. These controls enable videoselection at passenger monitors and audio volume adjustment at eachheadset, as well as enable video selection and audio volumeadjustment on the 15” monitor and the audio speakers, respectively.Each control unit display indicates which video or audio is selected.

Audio is supplied by the DVD players, which can function as CDplayers, the video cassette player, and the Airshow digital interfaceunit. The IFE system provides selectable audio and volume control ateach passenger seat location, and with the system offering priorityinterruption capability that enables passenger address tosimultaneously override all audio channels.

The Airshow digital interface unit generates text and graphic outputinformation obtained from the airplane´s long-range navigation system,air data system, and a CD-ROM. This information, selected throughthe Flight Deck Controller (FDC) in the cockpit, includes appropriatemaps and points of interest for the passengers and flight crew-members.


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MAIN COMPONENTS

AIRSHOW DIGITAL INTERFACE UNIT

The Airshow digital interface unit, powered by the entertainmentcabinet CB panel, is a computer that receives digital and analogsignals from the on-board aircraft avionics systems (FMS and ADC),and data from the magnastar connection. The Airshow digital interfaceunit is controlled through the MH Interface (MHI), which connects theAirshow with the MH data bus. The Airshow supplies the followingoutputs:− Geographic Maps of flight routes that show real-time airplane

position, previous flight path and planned route;− Multi-language location names and points of interest on the map;− Multi-language text pages of real-time flight information (e.g., ground

speed, altitude, temperature, distance and times);− Customized graphics, including customized logos and

announcements;− Audio briefings for safety and/or personal announcements.

There is a dedicated 6.5” LCD Airshow video output monitor thatenables the selection of any information before it is presented in thepassenger cabin.

FLIGHT DECK CONTROLLER

The flight deck controller, mounted in the cockpit, is responsible forselecting customer-request features in the Airshow digital interfaceunit, such as time to destination, Greenwich mean time, destinationairport, mode display, graphic display, languages and audioannouncements.

MH ENTERTAINMENT CONTROLLER

The MH entertainment (MHE) controller switches and distributesentertainment audio and video, and also provides chime to thepassenger cabin. It has eight video inputs switchable to the 24headphone outputs and four speaker outputs. The MHE is controlledthrough the MH data bus and supplied by the entertainment cabinet CBpanel.

CONTROL PANELS

There are several digital control panels connected to the MH data bus,which enable the audio and video selection at each seat location,and/or audio and video selection in the passenger cabin. The controlpanels are the Master Entertainment Control Panel, the Forward andAft VIP Control Panels, and the eight Passenger Control Panels.

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IN-FLIGHT ENTERTAINMENT PANEL

The In-Flight Entertainment (IFE) system should be activated ordeactivated if the IFE button is released on the IFE Panel.

1 - IN-FLIGHT ENTERTAINMENT BUTTON− Enables (released) or disables (pressed) the power supply to

the In-Flight Entertainment System.− A striped bar illuminates in the button to indicate that it is

pressed.

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IN-FLIGHT ENTERTAINMENT BUTTON ON THE IFE PANEL


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MASTER ENTERTAINMENT CONTROL PANEL


The master entertainment control panel is located on the entertainmentcabinet and it has switches to control audio and video functions of in-flight entertainment systems. It also brings a display that presentsentertainment-related information.The figure on the next page presents the functions of each controlbutton of the master entertainment control panel.

1 - DISPLAY− Shows the video and audio channels to be selected.

2 - MAP MODE SELECT− Selects the map modes of the Airshow.

3 - FWD AUDIO SELECT− Selects the audio for the forward passenger cabin.

4 - VOLUME UP KEY− Increases the volume on the audio speakers.

5 - VOLUME DOWN KEY− Decreases the volume on the audio speakers.

6 - AFT AUDIO SELECT− Selects the audio for the aft passenger cabin.

7 - AFT SPEAKER ON/OFF− When alternately pressed, turns ON or OFF the aft speakers.

8 - FWD SPEAKER ON/OFF− When alternately pressed, turns ON or OFF the forward

speakers.

9 - FWD MONITOR ON/OFF− When alternately pressed, turns ON or OFF the forward

monitor.

10 - FWD VIDEO SELECT− Selects the video source to be displayed on the forward monitor.

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FOR AIRPLANES FROM S/N 686 (INCLUSIVE) AND SUBSEQUENTAIRFRAMES.

The master entertainment control panel is located on the entertainmentcabinet and it has switches to control audio and video functions of in-flight entertainment systems. It also brings a display that presentsentertainment-related information.The figure on the next page presents the functions of each controlbutton of the master entertainment control panel.

1 - AIRSTAIR DOOR CLOSE

− Controls the main door to close.

2 - DOOR BLOCKED

− Controls the main door to block.

3 - AIRSTAIR LIGHTS ON/OFF

− Controls the stair lights on/off.

4 - STEP LIGHT ON/OFF

− Controls the step light on/off.

5 - AREA LIGHT ON/OFF

− Controls the door area light on/off.

6 - DISPLAY

− Monitors source selection

7 - SPEAKER SELECTION INDICATION

− Indicates if the speaker is selected.

8 - FORWARD VIDEO SELECTION INDICATION

− Indicates if the forward video is selected.

9 - AFT VIDEO SELECTION INDICATION

− Indicates if the aft video is selected.

10 - TEMPERATURE

− Displayed in degrees Fahrenheit.

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MASTER ENTERTAINMENT CONTROL PANEL(FOR AIRPLANES S/N 686 (INCLUSIVE) AND SUBSEQUENT

AIRFRAMES)


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11 - MAP DISPLAY MODE; (MAP, AUTO, LOGO, INFO)

− Selects the map modes of the Airshow.

12 - AFT LCD ON/OFF SWITCH

− When alternately pressed, turns ON or OFF the aft LCD.

13 - FORWARD LCD ON/OFF SWITCH

− When alternately pressed, turns ON or OFF the forward LCD.

14 - SPEAKER ON/OFF SWITCH

− When alternately pressed, turns ON or OFF the speaker.

15 - VOLUME UP KEY

− Increases the volume on the audio speakers.

16 - VOLUME DOWN KEY

− Decreases the volume on the audio speakers.

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SATCOM SYSTEM

The SATCOM system is a compact aeronautical system whichprovides one-channel data/voice/fax for direct satellite communicationin the INMARSAT satellite network.

The SATCOM is a single-channel AERO-M INMARSAT standarddevice, utilizing the spot beams of 3rd generation of INMARSATsatellite to provide a global continental communication. Through itssingle channel, the SATCOM supplies digital voice service at 4800 bpsand data service at 2400 bps.

The SATCOM system includes the following components:- Satellite Data Unit (SDU).- High-power/low noise amplifier.- SATCOM handset #1 and cradle.- SATCOM handset #2 and cradle (optional).- SATCOM antenna and NRS system.- SATCOM pots telephone adapter (optional).

The figure on the next page presents SATCOM system componentlocations.

For further information about the SATCOM system, refer toSection 2-18 in this volume.


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TELEPHONE SYSTEM

The MagnaStar C-2000 digital airborne telephone system transmitsand receives both voice and fax/modem type data to and from theaircraft, using an exclusive cellular-linked technology.

The telephone system has two air-ground voice/fax/datacommunications channels that can be used simultaneously. The digitalairborne telephone system operates over the “GenStar” satellitesystem, which offers continuous coverage throughout the UnitedStates of America, southern section of Canada and most section ofMexico.

The frequency range of operation to transmit is 894 to 896 MHz and toreceive is 849 to 851 MHz.

The telephone handsets contain noise-canceling microphones toovercome the aircraft noise environment, and the audio is digitized atthe handset to further ensure clear communication. Special functionkeys and a liquid crystal display (LCD) allow easy use of all options.The handset has an adjustable audio volume, telephone call styles,and selects various options. The LCD display presents usefulinformation and menu style selections.

There is a fax machine with fax/scanner/printer and copier functionsinstalled in the credenza.

The operation of the telephone system is similar to a cellular phone.When in normal operation, the handset works as a normal telephone.To place a call, it is necessary to pick up the handset and dial thedesired phone number with all prefix and country codes.

The figure on the next page presents the telephone handset and faxlocations.


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SECTION 2-07

FIRE PROTECTION

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-07-05 ..01Engine and APU Fire Protection System ........................... 2-07-10 ..01

Fire/Overheat Detection ................................................. 2-07-10 ..01Fire Extinguishing ........................................................... 2-07-10 ..04Controls and Indicators................................................... 2-07-10 ..06EICAS Messages ........................................................... 2-07-10 ..08

Lavatory and Galley Fire Protection System ...................... 2-07-15 ..01Lavatory and Galley Fire Extinguishing .......................... 2-07-15 ..01

Baggage Compartment Fire Protection System ................ 2-07-20 ..01Baggage Compartment Smoke Detection System......... 2-07-20 ..01Baggage Compartment Fire Extinguishing System........ 2-07-20 ..01EICAS Messages ........................................................... 2-07-20 ..02Controls and Indicators................................................... 2-07-20 ..04

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GENERALThe engine and APU fire protection system consists of fire/overheatdetection and a fire extinguishing system.

The fire protection for the lavatory compartment is equipped with adedicated smoke detection system and the lavatory waste container aswell as the forward galley are equipped with a fire extinguishingsystem.

In addition, The baggage compartment is also provided with a smokedetection system and with a fire extinguishing system.

The detection system provides visual and aural(except in the baggagecompartment) means of detecting a localized fire, smoke or generaloverheating. Also, a Monitoring circuitry is provided to continuouslycheck the fire and smoke detection systems and the fire extinguishingsystem and In case of failure, to signal the EICAS display.

Extinguisher bottles are installed to extinguish the fire in the airplane’sengines, APU, lavatory waste container, forward galley and baggagecompartment. Portable halon fire extinguishers installed at the frontand rear of the airplane can be used to extinguish small fires in thecockpit or main cabin area. A single water extinguisher is an additionaloption.


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ENGINE AND APU FIRE PROTECTION SYSTEM

FIRE/OVERHEAT DETECTION

The engines and the APU are protected against the occurrence of fireby means of fire detection and fire extinguishing systems.

Essential DC bus 1 powers the engine 1 fire detection system andessential DC bus 2 powers the engine 2 and the APU fire detectionsystem. Hot battery bus 1 and 2 power the engine fire extinguishingsystem, whereas the APU fire extinguishing system is powered byessential DC bus 2.

The fire/overheat detection system is provided with independentsensor tubes installed in the engines and APU. The sensor tubecontains a fixed volume of inert gas (Helium) and a gas-impregnated(Hydrogen) core material. The inert gas provides sensing ofoverheating. The core element provides sensing of localized fire orhigh-intensity heating. Overheating causes the sensor tube’s internalgas pressure to increase. This closes a switch on the fire/overheatingdetection system’s electrical circuit and activates the warning system.Localized fire or high-intensity heating increases the central core’s gasvolume, raising the sensor tube’s internal pressure, thus activating thealarm switch in the same manner as described above.

Manual resetting of the fire detection system is not available. Uponremoval of the fire or overheat condition, a reversible process takesplace, and the system automatically returns to the normal standbyoperation mode.

An integrity switch continuously monitors the sensor tube’s integrity.The integrity switch is held closed by the sensor’s internal pressure.Should this pressure be lost the integrity switch opens, generating asignal to indicate that the system is inoperative.

Upon detection of a fire/overheat signal in the engine or APU, theassociated handle (for the engines) illuminates, an aural warning isgenerated and a warning message is presented on the EICAS. Thevisual warning remains activated as long as the fire signal exists. Theaural warning may be canceled by pressing the master warning light.

In the case of failure of any fire detector, a caution message ispresented on the EICAS.

The APU fire detection system provides a signal to shut down the APUautomatically in case of fire warning during ground operation.


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FIRE OVERHEAT DETECTION SCHEMATIC


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FIRE EXTINGUISHING

Two fire extinguishing bottles for the engines and one for the APU areinstalled in the airplane’s tail cone.

The extinguishing agent discharge is accomplished by braking theextinguisher bottle’s seal through an electrically actuated cartridge inthe discharge valve.

Each engine fire extinguisher bottle contains two discharge valves, apressure gauge with a pressure switch and a fill/safety relief valve. Theengine bottles are cross-connected by two double check tees toprovide dual shot capability, thus one or both bottles can be dischargedinto one or the other engine. The double-check tee prevents theextinguishing agent of the remaining bottle from filling the emptiedbottle in case of a second shot of the system. The engine extinguisherbottles are discharged by pulling and rotating the Fire ExtinguishingHandle, which is located on the overhead panel.

CAUTION: DO NOT DISCHARGE THE SAME EXTINGUISHERBOTTLE TWICE. ACTUATING THE FIRE HANDLE INTOAN EMPTY BOTTLE MAY CAUSE STRUCTURALDAMAGE TO THE BOTTLE.

The APU bottle contains only one discharge valve, a pressure gaugewith a pressure switch, and a fill/safety relief valve. It provides singleshot capability for the APU. The APU extinguisher bottle is dischargedby pressing the APU Fire Extinguishing Button, located on theoverhead panel.

A caution message is presented on the EICAS should any bottle bedischarged or be inoperative for any reason (failed cartridge, loss ofpressure, or loss of power).

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ENGINE AND APU FIRE EXTINGUISHING SYSTEM SCHEMATIC


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ENGINE AND APU FIRE DETECTION/EXTINGUISHING SYSTEMPANEL

1 - ENGINE FIRE EXTINGUISHING HANDLE

− During normal flight conditions, the handle remains flush withthe panel.

− A red light illuminates inside the handle upon detection of fire oroverheating.

− When pulled, it closes the fuel, hydraulic, bleed air, and lip anti-icing shutoff valves of the associated engine.

− When rotated counterclockwise or clockwise, it respectivelydischarges extinguisher bottles A or B into the associatedengine.

2 - APU FIRE EXTINGUISHING BUTTON (guarded)

− When pressed, it closes the APU fuel shutoff valve anddischarges the APU fire extinguisher bottle.

3 - FIRE DETECTION SYSTEM TEST BUTTON

− When pressed and held for at least two seconds, it permits thefire detection system to be checked.

The EMB-135BJ is equipped with class “C” baggagecompartment, and the fire test is successfully completed if theconditions below occur simultaneously:− The following EICAS fire detection messages are displayed:

− Warning: APU FIRE, ENG 1 (2) FIRE, BAGG SMOKE

− Caution: APU FIREDET FAIL, E1 (2) FIREDET FAIL− Fire handles illuminate.− Baggage fire extinguishing button illuminates.− Baggage compartment fan deactivates.− WARNING/CAUTION lights flash.− Aural warning sounds.

NOTE: - On the ground, when pressed approximately for morethan 10 seconds, the APU is shut down, if it is running.

If it is necessary to repeat the test, wait at least 6seconds to press the test button again.

- If Fire Detection Test button is held for less than 2seconds the BAGG EXTG button may remainilluminated. In this case, repeat the test.

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ENGINE AND APU FIRE DETECTION/EXTINGUISHING PANEL


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EICAS MESSAGES

TYPE MESSAGE MEANINGWARNING APU FIRE Fire in the APU.

ENG1 (2) FIRE Fire in associated engine.E1 (2) FIREDET FAIL Associated engine fire

detection system failed.

CAUTIONAPU FIREDET FAIL APU fire detection system

failed.E1 (2) EXTBTLA INOP Associated bottle hasE1 (2) EXTBTLB INOP been discharged or isAPU EXTBTL INOP inoperative.

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LAVATORY AND GALLEY FIRE PROTECTIONSYSTEM

LAVATORY AND GALLEY FIRE EXTINGUISHING

The lavatory and forward galley fire extinguishing system consists oftwo autodischargeable fire extinguisher bottles. Each bottle is attachedto a waster container, one in the lavatory vanity assembly and the otherone in the forward galley.

The bottle is mounted with the discharging tubes extended into thewaste container. The end of each discharge tube has a tip which has amelting temperature of approximately 77°C (170°F). The tips will meltto discharge the agent totally into the waste container when a fireoverheat condition occurs.


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FORWARD GALLEY FIRE EXTINGUISHER BOTTLE

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LAVATORY FIRE EXTINGUISHER BOTTLE


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BAGGAGE COMPARTMENT FIRE PROTECTIONSYSTEM

BAGGAGE COMPARTMENT SMOKE DETECTIONSYSTEM

A smoke detection system is provided in the baggage compartment.The system consists of two smoke detection modules, one installed inthe compartment ceiling and the other one in the baggagecompartment aft bulkhead.A warning message is presented on the EICAS to indicate smokedetection inside the baggage compartment.The smoke sensor resumes the normal operation when the fire isextinguished, the smoke has been cleared and the smoke sensor isreset through the power reset button, located on each smoke detectionmodule.

NOTE: The Smoke Detector Modules are normally handled by themaintenance personnel.

In order to avoid inadvertent occurrences of smoke detectionmessages due to humidity in the smoke sensors, an integral heaterraises the temperature of the optical components of the smokedetector. In addition, a fan with brushless DC motor provides airperturbation in the vicinity of the detector.An alarm condition may also be triggered by high temperatureconditions in the absence of high smoke levels.

BAGGAGE COMPARTMENT FIRE EXTINGUISHINGSYSTEMTwo fire extinguishing bottles (High-rate Discharge Bottle and MeteringDischarge Bottle) are installed in the rear electronic compartment forfire baggage compartment protection.The High-rate Discharge Bottle is intended to fill the baggagecompartment instantaneously while the Metering Discharge Bottleprovides the concentration fire level extinguishing agent for at least 60minutes.Upon smoke detection inside the baggage, the smoke detectors senda signal to deactivate the baggage compartment fan (refer to Section2-2 - Equipment and Furnishings). Also, they illuminate the baggagefire extinguishing button on the Fire Detection/Extinguishing Panel andthe “DO NOT OPEN DOOR” warning on the lavatory aft bulkhead.


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EICAS MESSAGES


WARNING BAGG SMOKESmoke has been detectedinside the baggagecompartment.

CAUTION BAGG EXTBTL INOP

Any of the bottles havebeen discharged or areinoperative.

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BAGGAGE DETECTION/EXTINGUISHING PANEL

1 - BAGGAGE FIRE EXTINGUISHING BUTTON (guarded)

− When lit, button indicates that smoke was detected inside thebaggage compartment or that the fan has been deactivated.

− Button remains lit as long as there is smoke inside baggagecompartment.

− If the airplane is parked, the pilot must check the baggagecompartment, through the peephole, before pushing the BAGGEXTG SW.

− When pressed:− Discharges the baggage fire extinguishing bottles.− Deactivates the baggage compartment fan

NOTE: Fire extinguishing agent may activate the smoke detector.

2 - FIRE DETECTION SYSTEM TEST BUTTON

− Refer to ENGINE AND APU FIRE DETECTION/EXTINGUISHINGPANEL.

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BAGGAGE COMPARTMENT SMOKE DETECTOR

− Detectors are tested during Fire Detection System test.

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SMOKE INDICATION ON THE LAVATORY AFT BULKHEAD


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BAGGAGE COMPARTMENT SMOKE DETECTOR

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SECTION 2-08

FUEL

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-08-05 ..01Fuel Tanks ......................................................................... 2-08-05 ..02

Wing Fuel Tanks ............................................................ 2-08-05 ..02Auxiliary Fuel Tanks ....................................................... 2-08-05 ..02Fuel Tanks Capacities.................................................... 2-08-05 ..03Fuel Tanks Vent System ................................................ 2-08-05 ..04Engine and APU Fuel Distribution and Control .............. 2-08-05 ..05Auxiliary to Wing Tanks Fuel Transfer ........................... 2-08-05 ..08

EICAS Messages ............................................................... 2-08-05 ..11Controls and Indicators ...................................................... 2-08-05 ..14

Fuel System Panel ......................................................... 2-08-05 ..14MFD Bezel...................................................................... 2-08-05 ..16Fuel Page on MFD ......................................................... 2-08-05 ..17EICAS Indications........................................................... 2-08-05 ..20

Refueling and Defueling..................................................... 2-08-10 ..01Pressurized Refueling .................................................... 2-08-10 ..01Defueling ........................................................................ 2-08-10 ..03Refueling Panel .............................................................. 2-08-10 ..05

Fuel Measuring Stick.......................................................... 2-08-15 ..01Measuring Stick Tables .................................................. 2-08-15 ..03

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GENERALThe EMB-135BJ´s fuel system consists of two independent systems,one for each engine, interconnected by a crossfeed line. The fuelsystem ensures proper fuel supply to the engines and APU under alloperating conditions.

The system allows refueling and defueling operations to be performedeither by pressure or by gravity.

NOTE: The fuel weight values presented in this section of the manualare based on a fuel density of 0.809 kg/liter (6.751 lb/US Gal).


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FUEL TANKSThe airplane´s fuel storage system comprises two wing fuel tanks andtwo auxiliary fuel tank systems, that are composed of four auxiliary fueltanks.

WING FUEL TANKSThe wing fuel tanks are located at each wing and extend to the wingstub.

The fuel flows from the wing tip to the wing root by gravity. A collectorbox inside the wing stub keeps the electrical pump inlets submerged.To prevent pump cavitation, an ejector pump and flap valves ensurethat there is enough fuel in the collector box during wing-down anduncoordinated maneuvers.

AUXILIARY FUEL TANKS

The auxiliary fuel tank systems are composed of two forward fuel tanksand two aft fuel tanks. Each auxiliary fuel tank system has received adesignation, the left-hand system being designated as auxiliary fueltank system 1 while the right-hand system has been designated asauxiliary fuel tank system 2.

FORWARD AUXILIARY FUEL TANKS

The forward auxiliary fuel tanks are installed in the front section of thewing-to-fuselage fairing, divided in two totally separated compartments.Each one operates as an independent tank, the left-hand forward fuelcell being designated as forward auxiliary fuel tank 1 (FWD 1) while theright-hand fuel cell has been designated as forward auxiliary fuel tank 2(FWD 2).

The fuel flows from the forward section to the rear of both tanks bygravity. A collector box in the rearmost region of the tanks keeps theelectrical pumps inlets submerged. To prevent pump cavitation, flapvalves ensure that there is enough fuel in the collector box duringnose-down and uncoordinated maneuvers.

AFT AUXILIARY FUEL TANKS

The aft auxiliary fuel tanks are installed inside the rear area of theairplane fuselage, aft of the baggage compartment, in a pressurizedarea. Each one operates as an independent tank, designated as aftauxiliary fuel tank 1 (AFT 1) for the left-hand cell, and aft auxiliary fueltank 2 (AFT 2) for the right-hand cell.

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FUEL TANKS CAPACITIESFuel tank capacities are listed in the table below. The values refer tousable fuel:

FUEL TANKS CAPACITIES

Liters US Gal Kg lb

Wing Tank 1 3198 845 2587 5703

Wing Tank 2 3198 845 2587 5703

Wing TanksSub Total

6396 1690 5174 11406

AUXFWD 1 1056(1)

1112(2)279(1)

294(2)800(1)

900(2)1764(1)

1984(2)

FUS 1 AFT 1 822 217 660(1)

670(2)1455(1)

1477(2)

AUX FUS 1Sub Total

1878(1)

1934(2)496(1)

511(2)1460(1)

1570(2)3219(1)

3461(2)

AUXFWD 2 1056(1)

1112(2)279(1)

294(2)800(1)

900(2)1764(1)

1984(2)

FUS 2 AFT 2 822 217 660(1)

670(2)1455(1)

1477(2)

AUX FUS 2Sub Total

1878(1)

1934(2)496(1)

511(2)1460(1)

1570(2)3219(1)

3461(2)

TOTAL 10152(1)

10264(2)2682(1)

2712(2)8094(1)

8314(2)17844(1)

18328(2)

Conversion factors:− 3.785412 liter/US gallon or 0.264172 US gallon/liter− 1.245 liter/kg or 0.809 kg/liter− 0.4536 kg/lb or 2.2046 lb/kg

NOTE: 1) Max fuel capacity/weight allowed. Applicable to airplanesS/N up to 591 and Pre-Mod. SB 145LEG-28-0010.

2) Max fuel capacity/weight allowed. Applicable to airplanesS/N 625 and on or Post-Mod. SB 145LEG-28-0010.


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FUEL TANKS VENT SYSTEMThe wing, forward and aft auxiliary tanks are vented to theatmosphere.

The purpose of the vent system is to prevent damage to the fuel tanksdue to excessive buildup of positive or negative internal pressures andto provide ram air pressure within the tanks. This system also preventsfuel spillage during flight maneuvers and hard braking.

WING FUEL TANKS VENT SYSTEM

Each wing fuel tank is vented by a system that consists of a vent tankand a NACA air intake. The vent tank is located in the wing tip and isconnected to the fuel tank through two float valves. These valves allowat least one venting point to remain open between the vent tank andthe fuel tank under any flight condition. The vent tanks are connectedto outside air through a NACA air intake installed under the wing.

FORWARD AUXILIARY FUEL TANKS VENT SYSTEM

Each forward auxiliary fuel tank is vented by a system that consists oftwo float valves connected to a dedicated NACA air intake, installed onthe wing-to-fuselage fairing. These valves allow at least one ventingpoint, at each forward tank, to remain open between the NACA airintake and the fuel tank under any flight condition.

AFT AUXILIARY FUEL TANKS VENT SYSTEM

Each aft auxiliary fuel tank is vented by a system that consists of onefloat valve and one relief valve connected to a vent line that receivespressure from the cabin and is also connected to a port installed on thebottom of the fuselage.

Each aft auxiliary fuel tank is pressurized with air from the cabin by adedicated pressurization line. The float valve is installed at the end ofthis pressurization line to avoid the entry of fuel into it duringuncoordinated maneuvers.

The relief valve assures a maximum differential pressure between theinterior of the tank and the cabin, discharging the excess of pressure tothe atmosphere, through the port on the bottom of the fuselage.

An electrical shutoff vent valve, installed in the vent line, allows tankventilation when the airplane is on the ground and during pressurerefueling. This valve is also open when the air conditioning packs are

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turned off or the aircraft is not energized. During unpressurized flights,limited to 10000 ft ceiling, the shutoff vent valve is open and aft tankventilation is provided by the float valve and the vent port.

ENGINE AND APU FUEL DISTRIBUTION AND CONTROLThere are three electric pumps for each wing tank, which providepressurized fuel to the engines and APU. One pump is capable ofsupplying fuel for both engines, plus the APU, under all flight phases -except take-off and go-around. During take-off and go-around, at leasttwo electric pumps are required to supply fuel for both engines and theAPU.

Engine-driven fuel pumps will provide suction feed should the electricfuel pumps operation not be available.

Six knobs and one push-button located on the overhead fuel panelcontrol the electric pumps, crossfeed and fuel transfer operations. TwoPUMP PWR knobs energize/de-energize the electric pumps and twoPUMP SEL knobs select which wing fuel tanks pumps will be operatingto feed both engines and the APU. The remaining pumps will remainon standby. If fuel pressure drops below 6.5 psi, the remaining pumpsare automatically switched on and start cycling, until the pilot selectsone of them. The XFEED and the FUS TK XFER knobs control thecrossfeed and the fuel transfer operations, respectively. One FWDPUMP SEL push-button selects which forward auxiliary fuel tank pumpwill be operating during the fuel transfer operation.

Crossfeed operations should be performed in case of fuel imbalancebetween tanks or during low fuel level operations. The crossfeed knobacts over the crossfeed valve and over the electric pumps. Selectingthe knob to LOW1 or LOW2 will deenergize the pump associated tothe side with low level. The crossfeed valve will open connectingengine 1 and engine 2 fuel feed lines. The fully opened crossfeed valveposition is indicated on the EICAS by an advisory message. In case ofvalve failure, the EICAS displays a caution message.

NOTE: - Crossfeed selector knob must be OFF during takeoff andlanding.

- Crossfeed operation does not allow fuel transfer betweenwing tanks.


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Fuel for APU operation is normally supplied from the right side fuelsystem. Fuel from the left-hand system may be used by selecting thecrossfeed knob to LOW2. The APU fuel shutoff valve will close in thefollowing conditions:− APU master knob positioned to OFF.− By pressing the APU fuel shutoff button.− By pressing the APU fire extinguisher button.− Automatically, through the APU fire detection system in case of APU

fire on ground.

Sensors installed in the tanks and along the fuel lines provide signalsto indicate system failures and status. Such indications and messagesare shown on the MFD Fuel page as well as on the EICAS.

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FUEL SYSTEM SCHEMATIC


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AUXILIARY TO WING TANKS FUEL TRANSFERThe EMB-135 BJ is equipped with two independent fuel transfersystems that operate automatically to transfer fuel from the auxiliary tothe wing tanks.

System 1, designated FUS 1, is used to simultaneously transfer fuelfrom the FWD and AFT auxiliary fuel tanks 1 to the WING tanks.System 2, designated FUS 2, is used to simultaneously transfer fuelfrom the FWD and AFT auxiliary fuel tanks 2 to the WING tanks.

Both fuel transfer systems are optimized to be used under cruisingconditions, at altitudes of 10000 ft and higher. The transfer systemsare designed to perform the fuel transfer from the FWD tanks to theright WING tank 2 and from the AFT tanks to the left WING tank 1.

For airplanes S/N up to 591 and Pre-Mod. SB 145LEG-28-0010: across transfer line allows, when the AFT auxiliary fuel tank becomeempty, the remaining fuel in the FWD auxiliary fuel tank to besimultaneously transferred to the left and right WING tanks, thusavoiding a wing imbalance.

For airplanes S/N 625 and on or Post-Mod. SB 145LEG-28-0010: across transfer line allows the initial fuel quantity difference of the FWDtank to be transferred simultaneously to the left and right WING tanks,thus avoiding a wing imbalance. This difference is transferred at thebeginning of the fuel transfer operation, before the respective AFT tanktransfer starting.

NOTE: Fuel transfer operations during takeoff and landing isprohibited.

There are two electric pumps for each forward auxiliary fuel tank. Onepump transfers fuel to WING tank 2, while the other remains onstandby. On the AFT auxiliary fuel tanks, there is one electric pump pertank to transfer fuel to WING tank 1. When operating at altitudes of20000 ft and above, cabin pressurization can be used as a backup incase of pump failure.

Switching the FUS TK XFER knob to FUS 1 or FUS 2 allows selectionof the fuel transfer system that will be placed into operation. Fueltransfer starts automatically when the fuel quantity in at least oneWING tank achieves the starting level – 1900 kg (4189 lb). Theselected FWD auxiliary fuel tank transfer shutoff valve is commandedto open if WING tank 2 achieves the starting level and, on the selectedAFT auxiliary fuel tank, the transfer shutoff valve is commanded toopen if WING tank 1 achieves the starting level.

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For airplanes S/N 625 and on or Post-Mod. SB 145LEG-28-0010, thecross transfer valve is opened at the beginning of the transfer,whenever the FWD tank has more fuel than the AFT tank due to theprevious refueling operation. This logic allows transfer from the FWDtank until its quantity is equal to the respective AFT tank. After fuelquantities in the FWD and AFT tanks are equal, the cross transfervalve is commanded to close and the transfer from both AFT and FWDtanks occurs normally to the left and right WING tanks, respectively.

There is an indication on the MFD to show which valve is open andthat the fuel transfer system is operating. The respective transfershutoff valve of the FWD or AFT auxiliary fuel tanks is automaticallycommanded to close when fuel quantity in the corresponding WINGtank achieves the full condition. When the starting level is reachedagain, on any WING tank, the fuel transfer operation restarts.

For airplanes S/N up to 591 and Pre-Mod. SB 145LEG-28-0010 thissequence is automatically repeated until the selected AFT auxiliary fueltank is empty, closing the transfer shutoff valve, turning off the pumpand opening the cross transfer shutoff valve. With the cross transfershutoff valve open, the selected FWD auxiliary fuel tank transfershutoff valve is commanded to open when at least one WING tankreaches the starting level and the other WING tank is not in the fullcondition. The transfer shutoff valve is commanded to close when thefuel quantity on at least one WING tank corresponds to the fullcondition. This sequence is automatically repeated until the FWDauxiliary fuel tank is empty, closing the transference shutoff valves, thecross transfer shutoff valve and turning off the pump.

For airplanes S/N 625 and on or Post-Mod. SB 145LEG-28-0010 thissequence is automatically repeated until the selected auxiliary fuel tankis empty, closing the transfer shutoff valve, turning off the pump.

When the selected FWD and AFT auxiliary fuel tanks are empty, thefuel transfer system is automatically turned off. Fuel transfer from theother set of auxiliary fuel tanks will only be initiated when the pilotcommands such action by switching the FUS TK XFER knob to theapplicable system. In this case, fuel transfer will initiate even if the wingtanks have fuel quantity higher than 1900 kg (4189 lb).

In case of FWD auxiliary fuel tank pump failure, the pilot must selectthe remaining pump through the FWD PUMP SEL push buttons. If thefailed pump belongs to the AFT auxiliary fuel tank, and the flight isbeing conducted above 20000 ft, the system will automatically starttransferring fuel using the cabin air pressurization.


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AUXILIARY TO WING TANKS FUEL TRANSFERENCE SYSTEMSCHEMATIC

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EICAS MESSAGES

TYPE MESSAGE MEANINGFUEL 1(2) LO LEVEL The remaining fuel quantity

in the associated wing tankranges from 210 kg (463 lb)to 265 kg (584 lb), for leveledflight condition.

WARNING

FUEL XFER CRITICALorCHECK ACFT LOAD

Total fuel quantity in allauxiliary fuel tanks containsat least 800 kg (1764 lb) andat least one wing tank fuelquantity contains 1000 kg(2205 lb) or less.

NO TAKEOFF CONFIG Fuel transfer system isswitched on to the takeoffconfiguration setting.Message associated to auralwarning TAKEOFF FUEL.

E1 (2) FUEL LO PRESS Fuel pressure, in theassociate engine, is below6.5 psig (disabled whempressure reaches 9.5 psig).

FUEL TANK LO TEMP Fuel temperature inside wingtank 1 is equal or below–40°C.

CAUTION

FUEL XFEED FAIL Disagreement betweencrossfeed valve and knobposition.

(Continued)

FUEL XFEED MISCMD Crossfeed valve remainsopen after fuel imbalancecorrection - differencebetween wing tanks fuelquantities lower than 45 kg(100 lb) - or crew activatedthe wing fuel imbalancecorrection to the wing tankwith low level.

FUELING DOOR OPN Refueling panel access dooris open (inhibited duringtake-off and approach).


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TYPE MESSAGE MEANINGWING FUEL IMBALANCE Fuel quantity in one wing

tank differs by 363 kg(800 lb) from the other wingtank. Message is removedwhen the difference betweenwing tanks decreases tobelow 45 kg (100 lb).

APU FUEL LO PRESS Fuel pressure is below6.5 psi with APU operating(disabled whem pressurereaches 9.5 psig).

E1 (2) FUEL SOV INOP Associated shutoff valve isnot in the commandedposition.

CAUTION

APU FUEL SOV INOP Associated shutoff valve isnot in the commandedposition.

FUEL XFER 1(2) INOP Selected auxiliary fueltransfer system has failed.

(Continued) FUEL XFER CHECK Fuel transfer system is notactivated and fuel level hasreached 1850 kg (4079 lb) inat least one wing tank.

FUEL XFER OVFLOW Fuel transfer system was notinterrupted 30 seconds afterat least one wing tank hasreached 2450 kg.

FUSELAGE FUEL IMB Difference between fuelquantity in the forwardauxiliary fuel tank and fuelquantity in the aft auxiliaryfuel tanks of each auxiliaryfuel transfer system isoutside the approved limits(refer to AOM vol. 01 chapter01-36).

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CAUTION

FUEL TK VENT OPEN Vent valve is commanded toclose but it is not in the fullyclosed position.For airplanes with differentialpressure switch installed, themessage is also activatedwhen the aft relief valvefailures causing an aft tankoverpressure.

E1 (2) FUEL SOV CLSD Associated shutoff valve isclosed.

ADVISORY

APU FUEL SOV CLSD APU fuel shutoff valve isclosed. Message remains onfor 10 seconds after APUMaster Knob is set to off. Ifvalve has been commandedto close through APU FuelShutoff Button or APU FireExtinguishing Button themessage will remain on.

FUEL XFEED OPEN Crossfeed valve is open.


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FUEL SYSTEM PANEL1 - CROSSFEED SELECTOR KNOB

LOW1 - Opens the crossfeed valve and turns off wing tank 1'selectric fuel pump.

OFF - Closes the crossfeed valve.LOW2 - Opens the crossfeed valve and turns off wing tank 2's

electric fuel pump.

2 - WING TANK PUMP SELECTOR KNOB

− Selects which electric wing tank pump will be placed intooperation. The non-selected pumps will remain on standby.

3 - WING TANK PUMP POWER KNOB

ON - Energizes the associated fuel pump's circuit.OFF - Deenergizes the associated fuel pump's circuit.

4 - TRANSFER SYSTEM SELECTOR KNOB

FUS 1 - Energizes fuel transfer system 1.FUS 2 - Energizes fuel transfer system 2.OFF - Deenergizes the fuel transfer system.

5 - FORWARD AUXILIARY TANK PUMP SELECTOR PUSH-BUTTON

− Selects which electric forward auxiliary tank pump will be placedinto operation. The non-selected pump will remain on standby.

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FUEL SYSTEM PANEL


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MFD BEZEL

1 - FUEL SYSTEM AND RESET BUTTON− Pressing the FUEL button selects the fuel system page on MFD.

Pressing the button a second time resets the fuel used to zero.Fuel used must be reset individually on each MFD.

MFD BEZEL

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FUEL PAGE ON MFD

1 - DIGITAL WING TANK QUANTITY INDICATION(WINGTANK 1 AND WING TANK 2)− The digital wing fuel tank quantity indicator ranges from 0 to

7500 kg (0 to 15000 lb), with a digital resolution of 10 units,regardless of the unit being used (kg or lb), for WING TANK 1and WING TANK 2.

− Colors:− Green above 400 kg (880 lb).− Amber and boxed from 210 to 400 kg (450 to 880 lb).− Red and boxed below 210 kg (450 lb).

2 - DIGITAL FORWARD AUXILIARY TANK QUANTITYINDICATION (FWD 1 AND FWD 2)− The digital forward auxiliary fuel tank quantity indicator ranges

from 0 to 1000 kg (0 to 2200 lb), with a digital resolution of 10units, regardless of the unit being used (kg or lb), for FWD 1 andFWD 2.

− Color: green.

3 - DIGITAL TOTAL FUEL QUANTITY INDICATION− Indicates the total fuel quantity in all the tanks.

4 - ANALOGICAL FUEL QUANTITY INDICATION− Quantity is indicated by a vertical bar and a pointer. The colors

used are the same as for digital fuel quantity indication.

5 - WING TANK OPERATING PUMP INDICATION− This indicator displays A, B, C or OFF, depending on which

electric wing tank fuel pump is selected and whether it is on oroff.

− Color: green.− The indication blinks when the pump is cycling until the pilot

selects another pump.


MANUAL

2-08-05 Page

18 Code

01 REVISION 4

6 - DIGITAL TOTAL FUEL USED INDICATION− Indicates the total fuel used.− Color: Green under normal operation. Replaced by Amber

dashes (in flight) or amber zero (on ground) if any problem isverified.

7 - DIGITAL WING TANK FUEL TEMPERATURE INDICATION− Ranges from –60°C to +60°C with a resolution of 1°C.− Colors:

− Green above –40°C.− Amber and boxed below –40°C.

8 - DIGITAL AFT AUXILIARY TANK QUANTITY INDICATION(AFT 1 AND AFT 2)− The digital aft auxiliary fuel tank quantity indicator ranges from 0

to 1000 kg (0 to 2200 lb), with a digital resolution of 10 units,regardless of the unit being used (kg or lb), for FWD 1 and FWD2.

− Color: green.

9 - AFT AUXILIARY TANK FUEL TRANSFER INDICATION− This indicator displays A, P, T or OFF.− A indicates that electric fuel pump is operating.− P indicates that cabin air pressurization is being used to

accomplish fuel transfer.− T indicates that fuel transfer operation is in progress.− Color: green.

10 - FORWARD AUXILIARY TANK FUEL TRANSFER INDICATION− This indicator displays A, B, T or OFF.− A or B indicates which pump is in operation.− T indicates that fuel transfer is in progress.− Color: green.− The indication blinks when the pump is cycling until the pilot

selects another pump.

AIRPLANEOPERATIONS

MANUAL

FUEL


19 Code

01

MFD FUEL PAGE


MANUAL

2-08-05 Page

20 Code

01 REVISION 4

EICAS INDICATIONS

1 - FORWARD AUXILIARY TANK FUEL QUANTITY (FWD 1 AND FWD 2)− The digital forward auxiliary fuel tank quantity indicator ranges


− Color: green.

2 - AFT AUXILIARY TANK FUEL QUANTITY (AFT 1 AND ATF 2)− The digital forward auxiliary fuel tank quantity indicator ranges


− Color: green.

3 - WING TANK FUEL QUANTITY (WING TANK 1 AND WING TANK 2)− The digital wing fuel tank quantity indicator ranges from 0 to

7500 kg (0 to 15000 lb), with a digital resolution of 10 units,regardless of the unit being used (kg or lb), for WING TANK 1and WING TANK 2.

− Colors:− Green above 400 kg (880 lb).− Amber and boxed from 210 to 400 kg (450 to 880 lb).− Red and boxed below 210 kg (450 lb).

AIRPLANEOPERATIONS

MANUAL

FUEL


21 Code

01

EICAS INDICATIONS


MANUAL

2-08-05 Page

22 Code

01 REVISION 4


AIRPLANEOPERATIONS

MANUAL

FUEL


1 Code

01

REFUELING AND DEFUELINGRefueling and defueling operations may be performed either bypressure or by gravity. The refueling panel on the right side of thewing-to-fuselage fairing allows pressurized refueling/defuelingoperations. A gravity filler cap on the upper side of each wing allowsgravity filling. Dump valves and drain valves are used for gravitydefueling.

PRESSURIZED REFUELING

Pressurized refueling operations require an energized system. Thiscan be accomplished by either selecting the power selection switch toNORMAL or BATTERY.

The system allows pressure refueling operation for the wing tanks onlyor for the wing tanks followed by the auxiliaries, in a totally automaticsequence. Refueling is performed at a pressure range from 35 to 50psi.

NOTE: For airplanes with High Level Exceeding Indicationincorporated, a refueling automatic shutoff failure will beidentified by the HLEIS (High Level Exceeding IndicationSystem), that will sense, via one HLS (High Level Switch) ineach wing and auxiliary tanks, that the fuel level in the failedtank reached over the maximum quantity approved for thattank and will advise the operator by illuminating, on therefueling panel, the “STOP RFL” red indicating light. Theoperator shall interrupt immediately the refueling operation,relieving the pressure of the fueling source, avoiding anoverfilling and consequently a fuel spillage, and shall follow theprocedure to remove the extra fuel of the associated tank.


MANUAL

2-08-10 Page

2 Code

01 REVISION 4

PRESSURIZED REFUELING OF WING AND AUXILIARY TANKS

With the refueling system energized, WING + FUS position should beselected in the TANKS SELECTION switch. As fuel pressure is appliedon the adapter, the CLSD 1, CLSD 2, CLSD AFT and CLSD FWDlights pertaining to the WING, FUS 1 and FUS 2 refueling shutoffvalves will illuminate to indicate that these valves are closed. Selectingthe REFUELING switch to OPEN will open the wing tank refuelingshutoff valves, turning off CLSD 1 and CLSD 2 lights and startingrefueling operations. After wing tanks fuel level reaches 30 liters (7.9US gallons) below the tank full capacity, the respective shutoff valveswill close, turning the lights on again. The FUS 1 and FUS 2 refuelingshutoff valves will open, turning off the CLSD AFT and CLSD FWDlights, thus starting auxiliary tanks refueling. After the auxiliary fueltanks attain their selected quantities, FUS 1 and FUS 2 shutoff valveswill close, turning the lights on again. The refueling operation can bestopped at any time by commanding the REFUELING switch to theCLOSED position.

NOTE: FUS 1 and FUS 2 auxiliary fuel tanks systems cannot be filledwith more than:- 1460 kg each (2920 kg total for both systems) for airplanes

S/N up to 591 and Pre-Mod. SB 145LEG-28-0010.- 1570 kg each (3140 kg total for both systems) for airplanes

S/N 625 and on or Post-Mod. SB 145LEG-28-0010.

PRESSURIZED REFUELING OF WING TANKS ONLY

With the refueling system energized, WING position should beselected to the TANKS SELECTION switch. As fuel pressure is appliedon the adapter, the CLSD 1 and CLSD 2 lights pertaining to the WINGrefueling shutoff valves will illuminate to indicate that these valves areclosed. Selecting the REFUELING switch to the OPEN position willopen the wing tank refueling shutoff valves, turning off CLSD 1 andCLSD 2 lights and starting refueling operations. After wing tanks attaina fuel level of either 30 liters (7.9 US gallons) below the tank fullcapacity or the selected quantity, the respective shutoff valves willclose, turning the lights on again. The refueling operation can bestopped at any time by commanding the REFUELING switch to theCLOSED position.

AIRPLANEOPERATIONS

MANUAL

FUEL


3 Code

01

DEFUELINGThe system is designed to allow pressurized defueling of the wingtanks, using the same adapter as pressure refueling. This operationcan be performed using the electric fuel pumps installed in the wingtanks or by suction (4 psi max.) provided by an external source.Selecting the DEFUELING switch to OPEN will open the defuelingshutoff valve, thus allowing defueling operation. To defuel wing tank 1,the XFEED knob must be positioned to LOW2. The auxiliary tanksdefueling can be accomplished by transferring fuel to the wing tanksand then performing pressurized wing tanks defueling.

Gravity defueling of the wing tanks may be achieved by using thedump valves and opening the associated gravity refueling cap. Theremaining fuel can be totally drained through the drain valves. Defulingthe auxiliary tanks can be accomplished by transferring fuel to the wingtanks and then performing wing tank gravity defueling or by using theauxiliary tanks drain valves.

CAUTION: DO NOT RUN ELECTRIC PUMPS WITH FUELQUANTITY IN EACH TANK BELOW 30 LITERS (8 USGAL) OR 37 KG (54 LB).


MANUAL

2-08-10 Page

4 Code

01 REVISION 4

PRESSURE REFUELING/DEFUELING SYSTEM SCHEMATIC

AIRPLANEOPERATIONS

MANUAL

FUEL


5 Code

01

REFUELING PANEL

1 - POWER SELECTION SWITCH (guarded)NORMAL - Refueling system is energized by DC Bus 1.BATTERY - Refueling system is connected to Hot Bus 1.

2 - DEFUELING OPEN LIGHT (white)− Illuminates when the defueling shutoff valve is open.

3 - DEFUELING SWITCH (guarded)− Actuates the defueling shutoff valve to open or close.

4 - FUEL QUANTITY REMAINING INDICATOR− Displays fuel remaining as selected by the TK SEL/TEST

Switch.− The tank selected is identified by the letters L, R and T (L for the

left tank, R for the right tank and T for both sides).− The unit of measurement (kg or lb) is also displayed.− In case of failure, a FAIL inscription is displayed and the

refueling/defueling operation is interrupted.− The established accuracy of the EMB-135BJ airplane Fuel

Quantity Gauging System is:

− For the wing tanks: ± 2% of the provided indication plus± 0.75% of the total usable fuel, considering the approvedfuels and normal flight attitudes;

− For the auxiliary tanks: ± 4% of the provided indication plus± 2% of the total usable fuel, considering the approvedfuels and normal flight attitudes.

5 - TK SEL/TEST SWITCH (spring loaded to center position)TEST - Initiates indicator built-in and probes conditions test. All

light segments illuminate and a failure code is presented,if a failure is detected.

TK SEL - Selects which fuel quantity is going to be displayed in theupper display. When the indicator is energized, the totalwings fuel quantity is shown. Sequentially actuating theswitch will select wing tank 1, wing tank 2, FUS 1 tanksand FUS 2 tanks fuel quantity.

6 - INCR/DECRT SWITCH (spring loaded to center position)− Increases or decreases fuel quantity selected value.− If moved from the neutral position during refueling, it interrupts

the operation. The refueling operation will be restored 4 secondsafter switch release.


MANUAL

2-08-10 Page

6 Code

01 REVISION 4

7 - FUEL QUANTITY SELECTED INDICATOR− Displays total fuel quantity set through the Quantity Select

Switch.− If a FAIL inscription is displayed in the FUEL QTY REMAINING,

this indicator shows the failure code.− The indicator displays zero as the refueling compartment door is

opened.

8 - REFUELING SWITCH (guarded)− Acts over the pilot valves to open or close the refueling shutoff

valves.

9 - REFUELING CLOSED LIGHTS (white)− Illuminate when the associated refueling line is pressurized and

the associated shutoff valve is closed.

10 - AUXILIARY TANKS ISOLATION VALVES LIGHTS (white)− Illuminate when the associated isolation valve is closed.

11 - WING + FUS/WING ONLY SWITCHWING + FUS - Selects the wing tanks followed by the auxiliary

tanks to be refueled.WING ONLY - Selects only the wing tanks to be refueled.

12 - TANK INDICATION LIGHT (white)− Illuminates when the associated tank is selected.

13 - STOP REFUELING LIGHTS (red)− Illuminate when fuel level in the failed tank reached over the

maximum quantity approved for that tank (For airplanes withHigh Level Exceeding Indication incorporated).

14 - FUS 1/FUS 2/WING SWITCHFUS 1 - Selects FUS 1 tanks, showing its data in the refueling

panel.FUS 2 - Selects FUS 2 tanks, showing its data in the refueling

panel.WING - Selects the wing tanks, showing its data in the refueling

panel.

NOTE: When the refueling panel door is closed, all the switches arepushed to the NORMAL or CLOSED position by their guards.

AIRPLANEOPERATIONS

MANUAL

FUEL


7 Code

01

REFUELING PANEL

4

13


MANUAL

2-08-10 Page

8 Code

01 REVISION 4


AIRPLANEOPERATIONS

MANUAL

FUEL


1 Code

01

FUEL MEASURING STICKTwo direct quantity measuring sticks under each wing and one in thewing stub permit to separately check the fuel quantity in each wingtank. Each measuring stick provides visual indication of the total fuelquantity of the associated side.

The table below provides minimum and maximum stick values:

STICK FUEL QUANTITYPOSITION LITERS US GAL

Wing Stub Min 7 2

Tank Max 783 207

Internal Min 781 206

Point Max 2168 573

External Min 2163 571

Point Max 2791 737


MANUAL

2-08-15 Page

2 Code


MEASURING STICK POINTS

AIRPLANEOPERATIONS

MANUAL

FUEL


3 Code

01

MEASURING STICK TABLES

To determine the fuel quantity using the direct quantity measuringsticks, the airplane must be laterally leveled with roll and pitch anglesbetween 0° to -2°. After refueling the airplane, start with the externalmeasuring stick, closest to the wing tip. On EMB-135 BJ airplanes,between 2163 and 2791 liters (571 and 737 US gal), the externalmeasuring stick provides a correct fuel quantity indication. Above2791 liters (737 US gal), it is not possible to measure the fuel quantitythrough the sticks. If the external measuring stick provides a zeroindication, use the internal measuring stick to obtain the fuel quantity.Between 781 and 2168 liters (206 and 573 US gal) the internalmeasuring stick provides a correct fuel quantity indication. If theinternal measuring stick provides a zero indication, use the wing stubstick to obtain the fuel quantity. It is not possible to measure the fuellevel through the measuring sticks if it is below 7 liters (2 US gal)either.

Enter the following measuring stick tables with the value read on thestick to obtain the fuel quantity (liters or US gallons). To find the fuelmass in kg (lb) multiply the volume in liters (US gal) by the actual fueldensity in kg/l (lb/US gal).

NOTE: Do not add measuring sticks values.


MANUAL

2-08-15 Page

4 Code


FUEL QUANTITY

STICKINTERNAL STICK EXTERNAL STICK WING STUB STICK

IND LITERS US GAL LITERS US GAL LITERS US GAL

0.1 781 206 2163 571 7 20.2 794 210 2184 577 14 40.3 805 213 2203 582 22 60.4 817 216 2220 586 29 80.5 827 219 2235 590 36 100.6 838 221 2249 594 43 110.7 848 224 2263 598 50 130.8 858 227 2276 601 56 150.9 868 229 2288 605 63 171.0 878 232 2301 608 69 181.1 888 235 2314 611 74 201.2 898 237 2327 615 80 211.3 907 240 2340 618 85 231.4 917 242 2354 622 90 241.5 927 245 2368 626 95 251.6 937 248 2383 630 100 261.7 947 250 2398 633 104 281.8 957 253 2413 638 109 291.9 968 256 2429 642 113 302.0 978 258 2445 646 116 312.1 989 261 2461 650 120 322.2 1000 264 2477 654 123 332.3 1011 267 2493 659 129 342.4 1022 270 2508 663 136 362.5 1034 273 2524 667 140 372.6 1045 276 2539 671 144 382.7 1057 279 2553 674 148 39

MEASURING STICK TABLES (SHEET 1 OF 4)

AIRPLANEOPERATIONS

MANUAL

FUEL


5 Code

01

FUEL QUANTITY

STICKINTERNAL STICK EXTERNAL STICK WING STUB STICK

IND LITERS US GAL LITERS US GAL LITERS US GAL

2.8 1069 282 2567 678 155 412.9 1081 286 2580 682 164 433.0 1093 289 2593 685 173 463.1 1105 292 2604 688 181 483.2 1118 295 2615 691 188 503.3 1130 299 2625 693 195 513.4 1143 302 2634 696 201 533.5 1156 305 2643 698 207 553.6 1169 309 2651 700 213 563.7 1182 312 2658 702 218 583.8 1194 316 2666 704 224 593.9 1208 319 2673 706 229 604.0 1221 322 2680 708 234 624.1 1234 326 2687 710 239 634.2 1247 329 2696 712 244 644.3 1260 333 2706 715 249 664.4 1273 336 2717 718 254 674.5 1286 340 2731 721 259 684.6 1300 343 2747 726 265 704.7 1313 347 2767 731 270 714.8 1326 350 2791 737 276 73



MANUAL

2-08-15 Page

6 Code


FUEL QUANTITY

STICK INTERNAL STICK WING STUB STICK

INDICATION LITERS US GAL LITERS US GAL

4.9 1339 354 281 745.0 1352 357 287 765.1 1365 361 293 785.2 1378 364 300 795.3 1391 368 306 815.4 1404 371 313 835.5 1418 374 320 855.6 1431 378 327 865.7 1444 381 334 885.8 1457 385 342 905.9 1470 388 350 926.0 1483 392 357 946.1 1496 395 365 976.2 1509 399 374 996.3 1522 402 382 1016.4 1535 405 390 1036.5 1548 409 399 1056.6 1561 412 408 1086.7 1574 416 416 1106.8 1588 419 425 1126.9 1601 423 434 1157.0 1615 427 443 1177.1 1629 430 452 1197.2 1642 434 461 1227.3 1656 438 470 1247.4 1670 441 479 1267.5 1685 445 488 1297.6 1699 449 497 131


AIRPLANEOPERATIONS

MANUAL

FUEL


7 Code

01

FUEL QUANTITY

STICK INTERNAL STICK WING STUB STICK

INDICATION LITERS US GAL LITERS US GAL

7.7 1714 453 505 1347.8 1729 457 514 1367.9 1744 461 523 1388.0 1759 465 532 1418.1 1774 469 541 1438.2 1790 473 550 1458.3 1806 477 558 1478.4 1822 481 567 1508.5 1838 486 576 1528.6 1855 490 585 1548.7 1872 494 593 1578.8 1889 499 602 1598.9 1906 503 611 1619.0 1923 508 620 1649.1 1941 513 629 1669.2 1958 517 639 1699.3 1976 522 648 1719.4 1994 527 658 1749.5 2012 531 668 1769.6 2030 536 679 1799.7 2048 541 689 1829.8 2065 546 701 1859.9 2083 550 713 18810.0 2101 555 725 19210.1 2118 560 738 19510.2 2135 564 752 19910.3 2152 568 767 20310.4 2168 573 783 207



MANUAL

2-08-15 Page

8 Code



AIRPLANEOPERATIONS

MANUAL

HYDRAULIC

JUNE 20, 2002 2-11-00 Page

1 Code

01

SECTION 2-11

HYDRAULIC

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-11-05 ..01System Description ............................................................ 2-11-05 ..02EICAS Messages ............................................................... 2-11-05 ..05Controls and Indicators ...................................................... 2-11-05 ..06

HYDRAULIC AIRPLANEOPERATIONS

MANUAL

2-11-00 Page

2 Code



AIRPLANEOPERATIONS

MANUAL

HYDRAULIC

JUNE 20, 2002 2-11-05 Page

1 Code

01

GENERALThe airplane is equipped with two independent hydraulic systems, eachpowered by one engine driven-pump and one electric motor-drivenpump. Both hydraulic systems are identical, except for the serviceseach system provides and a priority valve installed in the hydraulicsystem 1.

There are ground connections for refilling and ground tests purposes.Indications of hydraulic system parameters are provided on the MFDand EICAS displays.

The services provided by each hydraulic system are presented below:

SYSTEM HYDRAULIC POWER SUPPLY

Ailerons SYSTEM 1 and 2

Rudder SYSTEM 1 and 2

Landing Gear SYSTEM 1

Main door SYSTEM 1

Steering SYSTEM 1

Brakes (Outboard Wheels) SYSTEM 1

Brakes (Inboard Wheels) SYSTEM 2

Emergency/Parking Brake SYSTEM 2

Thrust Reverser 1 SYSTEM 1

Thrust Reverser 2 SYSTEM 2

Outboard Spoilers SYSTEM 2

Inboard Spoilers SYSTEM 1


MANUAL

2-11-05 Page

2 Code

01 JUNE 20, 2002

SYSTEM DESCRIPTIONEach hydraulic system consists of a hydraulic fluid reservoir, amanifold, one engine-driven pump, one electric motor-driven pump,one shutoff valve, one accumulator and a priority valve installed in thehydraulic system 1.

RESERVOIRThe hydraulic fluid stored in the reservoir is pressurized, to avoid pumpcavitation. This pressurization function is performed by fluid drainedfrom the pressure line. The reservoir is equipped with a quantityindicator which transmits information to the MFD and EICAS displaysfor indication and warning purposes. A thermal switch is responsiblefor the high temperature message, if the fluid temperature increasesabove 90°C.

SHUTOFF VALVEA shutoff valve is installed between the reservoir and the engine-drivenpump. It cuts the hydraulic fluid supply to the engine-driven pump, ifthere is a fire on the related engine or in case of hydraulic fluidoverheat. This valve may be closed either through the engine fireextinguishing handle or through a dedicated button on the overheadpanel.

ENGINE-DRIVEN PUMPThe engine-driven pump provides continuous fluid flow at 3000 psi foroperation of the various airplane hydraulically-powered systems. Thepump is connected to the engine accessory gearbox and, as long asengine is running, it generates hydraulic pressure. During engine start,the fluid remaining in the suction line is sufficient to avoid pumpcavitation and provide reservoir pressurization.

ELECTRIC MOTOR-DRIVEN PUMPThe electric motor-driven pump has the same connections as theengine-driven pump, but has a lower flow capacity. The pump normallyoperates in the automatic setting mode, turning on when theassociated hydraulic pressure drops below 1600 psi or the associatedengine N2 drops below 56.4%.If the pump starts operating in the automatic mode, it will be turned offafter the pressure or N2 are reestablished to normal values. Theelectric pump may be turned on at pilot command, through the selectorknob on the overhead panel, furnishing continuous fluid flow at 2900psi.

AIRPLANEOPERATIONS

MANUAL

HYDRAULIC

JUNE 20, 2002 2-11-05 Page

3 Code

01

HYDRAULIC SYSTEM SCHEMATIC


MANUAL

2-11-05 Page

4 Code

01 JUNE 20, 2002

MANIFOLDThe manifold provides the following functions:

- Fluid filtering (pressure and return lines).- Overpressure relief (main and electrical pumps).- Pressure indications (main and electrical pumps).

Fluid leaving the pump flows to the manifold, where it is filtered andthen routed to the airplane systems. Inside the manifold, a check valveprevents the fluid from returning to the pump, while a relief valvediverts the excess fluid to the return line. The return line is supplied bythe fluid coming from the airplane systems, fluid drained from thepump, fluid from the relief valve, and fluid refilled by the maintenancepersonnel. Under any situation the fluid is filtered and returned to thereservoir. The manifold incorporates two pressure switches to detectlow hydraulic pressure, and a pressure transducer to indicate systempressure. Signals from the pressure switches and pressure transducerare sent to the MFD and EICAS displays.

PRIORITY VALVEThe hydraulic system 1 incorporates a priority valve. If the system ispowered by the electric motor-driven pump and the landing gear iscommanded to retract, the valve will provide minimum flow to thelanding gear system and give priority to the flight control services. Inthis case, the landing gear will operate through the accumulatorpressure.

ACCUMULATOREach hydraulic system has one accumulator. The function of theaccumulator is to keep the surges of the hydraulic pumps at aminimum, and to keep a 3000 psi pressure available for operation ofthe landing gear and main door (system 1) or operation of theemergency parking brake (system 2).

AIRPLANEOPERATIONS

MANUAL

HYDRAULIC

JUNE 20, 2002 2-11-05 Page

5 Code

01


CAUTION

HYD SYS 1 (2) FAIL Associated hydraulic systemis not pressurized (inhibitedwhen the airplane is on theground, engine is shut downand parking brake isapplied).

HYD SYS 1 (2) OVHT Associated hydraulic systemfluid temperature is above90°C.

E1 (2) HYD PUMP FAIL Engine-driven pump is notgenerating pressure withassociated engine running.

E1 (2) HYDSOV CLSD Associated hydraulic shutoffvalve is closed.

ADVISORY HYD1 (2) LO QTY Fluid level in the associated

reservoir is below one liter.Report to the maintenancepersonnel if the hydraulicreservoir operates empty.

HYD PUMP SELEC OF Associated electric pumpselected OFF with theparking brake released.


MANUAL

2-11-05 Page

6 Code

01 JUNE 20, 2002


HYDRAULIC SYSTEM PANEL

1- ENGINE PUMP SHUTOFF BUTTON (guarded)− Closes (pressed) or opens (released) the associated engine

pump shutoff valve.

− A striped bar illuminates in the button to indicate that it ispressed.

2- ELECTRIC HYDRAULIC PUMP CONTROL KNOBOFF - Associated pump is turned off.AUTO - Associated pump is kept in standby mode, ready to operate

if the engine-driven pump outlet pressure drops below 1600psi or the associated engine N2 drops below 56.4%.

ON - Associated pump is turned on.

AIRPLANEOPERATIONS

MANUAL

HYDRAULIC

JUNE 20, 2002 2-11-05 Page

7 Code

01

HYDRAULIC SYSTEM PANEL


MANUAL

2-11-05 Page

8 Code

01 JUNE 20, 2002

HYDRAULIC PAGE ON MFD

1- FLUID QUANTITY INDICATION− Ranges from zero to maximum hydraulic fluid quantity.

− Scale (horizontal line) and pointer:

− green when greater than 1 liter.

− amber when equal to or less than 1 liter.

− Pointer disappears if data is invalid.

2- PRESSURE INDICATION− Ranges from 0 to 4000 psi, with a resolution of 100 psi.

− Digits:

− green from 1300 to 3300 psi.

− amber and boxed below 1300 and above 3300 psi.

− Digits are replaced by amber dashes if data is invalid.

3- ELECTRIC PUMP STATUS− Indicated by the green label ON or OFF.

HYDRAULIC PAGE ON MFD

AIRPLANEOPERATIONS

MANUAL

LANDING GEARAND BRAKES


1 Code

01

SECTION 2-12

LANDING GEAR AND BRAKES

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-12-05 ..01Air/Ground Indication System ............................................ 2-12-05 ..03Landing Gear Operation..................................................... 2-12-05 ..04

Landing Gear Retraction ................................................ 2-12-05 ..04Landing Gear Extension ................................................. 2-12-05 ..06Landing Gear Warning ................................................... 2-12-05 ..08EICAS Messages ........................................................... 2-12-05 ..09Controls and Indicators................................................... 2-12-05 ..09

Brake System..................................................................... 2-12-10 ..01Normal Brake System .................................................... 2-12-10 ..02Emergency/Parking Brake System................................. 2-12-10 ..08EICAS Messages ........................................................... 2-12-10 . 10Controls and Indicators................................................... 2-12-10 . 10

Nose Wheel Steering System ............................................ 2-12-15 ..01EICAS Messages ........................................................... 2-12-15 ..02Controls and Indicators................................................... 2-12-15 ..04EMB-135 BJ Minimum Turning Radii ............................. 2-12-15 ..07


AIRPLANEOPERATIONS

MANUAL

2-12-00 Page

2 Code



AIRPLANEOPERATIONS

MANUAL



1 Code

01

GENERALThe EMB-135 BJ landing gear incorporates braking and steeringcapabilities. The extension/retraction, steering and braking functionsare hydraulically assisted, electronically controlled and electronicallymonitored. EICAS indications and messages alert crew to systemstatus and failures. Each landing gear is equipped with alternatemeans of actuation in case of normal actuation system failure.


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

2 Code



AIRPLANEOPERATIONS

MANUAL



3 Code

01

AIR/GROUND INDICATION SYSTEMAir/ground indication is determined by a system that detects landinggear shock absorber compression and relays information to thelanding gear electronic unit for gear control. The system consists offive weight-on-wheel proximity switches. Two of them are installed oneach main landing gear leg and one on the nose landing gear leg.

The Landing Gear Electronic Unit (LGEU) processes the main landinggear proximity switches’ signals information in four independentchannels and controls various equipment operations. Logic processingincludes the position signal and its validity. If all proximity switchsignals are valid, four signals are processed to assure that at leastthree signals indicate identical status for releasing the air/ground signaloutput.

Should one proximity switch signal be invalid, the logic will process theremaining three signals so that at least two indicate the same status. Ifa second proximity switch is invalid, the two remaining signals areprocessed only if both send the same signal. Disagreement betweenthese two remaining proximity switches causes the Landing GearElectronic Unit to de-energize the channels and provide a defaultoutput signal.

The nose landing gear proximity switch signal is sent only to the thrustreverser logic (if installed) and steering control.


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

4 Code


LANDING GEAR OPERATIONLanding gear retraction and extension are powered by the hydraulicsystem 1. An accumulator prevents pressure fluctuations and assistsgear retraction after takeoff. The main landing gear legs retractinboard, while the nose landing gear retracts forward. Each main gearleg is mechanically linked to its respective door, which remains openwhen the gear is down. The doors close automatically when the mainlanding gear is retracted. The nose landing gear doors are hydraulicallyactuated and operate in sequence with the nose gear.

Gear retraction and extension are electrically commanded. If normalextension fails, the landing gear can be extended through an electricaloverride system. If the electrical override is not available, a free-fallsystem allows gear extension. Gear position is indicated on the EICASdisplay.

LANDING GEAR RETRACTION

Landing gear retraction is commanded through the Landing GearLever, installed on the main panel. Positioning the lever to the UPposition signals the LGEU to command the Nose Gear Door SolenoidValve and the Landing Gear Electrovalve. This allows pressure fromthe hydraulic system 1 to simultaneously reach landing gear and downunlock actuators. All gear legs are then retracted into their respectivewheel wells.

The LGEU logic only allows the nose gear doors to close after the noselanding gear is locked in the UP position. When the uplock boxes areactuated, the proximity switches signal the LGEU that the gear is upand locked and that the Landing Gear Electrovalve may bedeenergized. Nose landing gear door actuators are kept pressurized,but the gear actuator lines are connected to the return.

AIRPLANEOPERATIONS

MANUAL



5 Code

01

LANDING GEAR SCHEMATIC


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

6 Code


To preclude an inadvertent retraction command while on the ground,the air/ground system provides a signal to a solenoid inside theLanding Gear Lever. This locks the lever and prevents movementtowards the UP position. For emergency purposes only, a lock releasebutton is provided beside the lever, allowing this protection to beoverriden.

LANDING GEAR EXTENSION

NORMAL EXTENSION

Positioning the Landing Gear Lever to the DOWN position signals theLGEU to command the Landing Gear Electrovalve and the Nose GearDoors Solenoid Valve. This allow pressure from the hydraulic system 1to simultaneously reach the landing gear and door actuators, and alsothe up unlock actuators.

When the gear legs reach the down position, the down lock boxes areactuated. The proximity switches signal the LGEU that the gear isdown and locked and that the Landing Gear Electrovalve may bede-energized.

ELECTRICAL OVERRIDE EXTENSION

The Electrical Override system is used to extend the landing shouldthere occur a normal landing gear extension failure. This systembypasses the LGEU and actuates directly the Landing GearElectrovalve and the Nose Gear Doors Solenoid Valve. The controlswitch is installed inside the free-fall lever compartment, on the floor,beside the copilot’s seat. Extension through override is made in steps,first opening the doors and then extending the gear. When extension iscompleted, selecting the override switch to normal positiondeenergizes the Landing Gear Electrovalve and depressurizes all lines.The switch is safeguarded, being in the non-actuated positionwhenever the compartment door is closed.

AIRPLANEOPERATIONS

MANUAL



7 Code

01

FREE-FALL EXTENSION

Free-Fall extension is available in case of failure of both normalextension and electrical override extension. Actuation of free-falllanding gear extension is performed by pulling up the lever installedinside the free-fall lever compartment, on the floor, beside the copilot’sseat.This mechanically actuates the Free-Fall Selector Valve and unlocksthe three landing gear legs uplocks. The Free-Fall Selector Valveisolates the hydraulic system pressure and connects the landing gearsystem hydraulic lines to the return. With the system unpressurizedand the uplocks deactivated, all gear legs fall by gravity until they reachtheir downlock devices. If one main gear does not lock down, increasethe aerodynamic drag by side slipping the aircraft to help lock theaffected leg.Once actuated, the free-fall lever remains locked in the vertical positionuntil mechanically released.


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

8 Code


LANDING GEAR WARNING

A LANDING GEAR voice message is provided to alert pilots any timethe airplane is in a landing configuration and the gear legs are notlocked down. The warning may be activated under one of threeconditions:

1. Radio Altitude below 1200 ft, Flap Selector Lever set below 22°,one thrust lever set below 59° and the other thrust lever set below45° (or the associated engine inoperative).

NOTE: In case of Radio Altimeter loss, the message may beactivated at any altitude, but may be canceled through theLanding Gear Warning Cutout Button.

2. Radio Altitude below 1200 ft, Flap Selector Lever between 22° and45°, one thrust lever set below 59° and the other thrust lever setbelow 45° (or the associated engine inoperative).

NOTE: - The Voice message cannot be canceled.- In case of Radio Altimeter loss, the message may be

activated at any altitude.

3. Flap Selector Lever set at 45°.

NOTE: The Voice message cannot be canceled.

AIRPLANEOPERATIONS

MANUAL


JANUARY 17, 2003 2-12-05 Page

9 Code

01

EICAS MESSAGES


WARNING

LG/LEVER DISAGREE After 20 seconds of gearcommand, at least onelanding gear is not in theselected position.

CAUTION

LG AIR/GND FAIL LGEU failure or failure of twoweight-on-wheel proximityswitches.

NLG UP/DOOR OPN(if applicable)

Nose LG is locked up andnose LG door is open.


LANDING GEAR CONTROL BOX

1 - LANDING GEAR LEVERUP - Selects landing gear retraction.DOWN - Selects landing gear extension.

2 - DOWNLOCK RELEASE BUTTON− − Mechanically releases the lever lock, allowing the landing gear

lever to be moved to the UP position when on the ground or incase it cannot be moved to the UP position after takeoff.

LANDING GEAR CONTROL BOX


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

10 Code


FREE-FALL LEVER COMPARTMENT

1 - FREE-FALL LEVER− When pulled up, depressurizes the landing gear hydraulic line

and releases all gear uplocks.− The lever is kept at the actuated position by a mechanical lock.

2 - FREE-FALL LEVER UNLOCK BUTTON− When pressed, unlocks the free-fall lever, allowing it to be

returned to the normal position, thus restoring the hydraulicoperation of the landing gear.

3 - ELECTRICAL OVERRIDE SWITCH (guarded)NORMAL - Landing gear retraction and extension are automatically

performed and controlled by the Landing GearElectronic Unit.

DOORS - Opens the nose landing gear doors.

GEAR/DOORS - Extends the landing gear.

FREE-FALL LEVER COMPARTMENT

AIRPLANEOPERATIONS

MANUAL



11 Code

01

LANDING GEAR WARNING CUTOUT BUTTON (guarded)

− When pressed, this button cancels the landing gear warning voicemessage if the Radio Altimeter is inoperative.

− An amber indication bar illuminates inside the button and remainsilluminated to indicate that a cancel action was performed.

− The amber indication bar extinguishes if the Thrust Levers areadvanced or Flap Selector Lever is set above 22° or landing gear isdown and locked.

LANDING GEAR WARNING CUTOUT BUTTON


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

12 Code


GLARESHIELD PANEL

1 - NOSE LANDING GEAR DOORS INDICATION LIGHT (if installed)− Illuminates to indicate that the nose landing gear is locked in the

retracted position and at least one door is not closed.

GLARESHIELD PANEL

EICAS INDICATIONS

1 - LANDING GEAR POSITION− Position is indicated by three boxes, one for each gear.

− Landing gear down and locked is indicated by a green DN labelinside a green box.

− Landing gear in transit is indicated when the box is cross-hatched in amber and black.

− Landing gear up and locked is indicated by a white UP labelinside a white box.

− Landing gear lever disagreement (landing gear is not in theselected position after 20 seconds) is indicated by a box cross-hatched in red and back or by a red label (UP or DN) inside ared box.

− Indication of landing gear downlocked is also presented on theRMU through the green LG DOWN LOCKED legend.

AIRPLANEOPERATIONS

MANUAL



13 Code

01

LANDING GEAR POSITION INDICATION ON EICAS


AIRPLANEOPERATIONS

MANUAL

2-12-05 Page

14 Code


LANDING GEAR INDICATIONS

AIRPLANEOPERATIONS

MANUAL



1 Code

01

BRAKE SYSTEMThe braking system consists of the normal brake system,emergency/parking brake system, and gear-retracting-in-flight braking.The normal brake system is supplied by hydraulic systems 1 and 2. Itis electronically commanded and monitored. The emergency/parkingbrake system is supplied only by hydraulic system 2 and ismechanically actuated. Normal braking is controlled by the pedals.Emergency braking is controlled by the emergency/parking brakehandle. Gear-retracting-in-flight braking is controlled by both hydraulicsystems and by a mechanical stop within the nose gear wheel well.This braking is electronically commanded and monitored.

Braking through the pedals incorporates some protections notavailable when using the emergency brake handle. Brake temperatureis shown on the MFD Hydraulic Page.


AIRPLANEOPERATIONS

MANUAL

2-12-10 Page

2 Code


NORMAL BRAKE SYSTEM

Normal brake system is operated by rudder pedal inputs. The brakesare powered by two independent hydraulic systems. It is controlled andmonitored by the Brake Control Unit (BCU). The BCU receives signalsfrom the pedal position transducers and commands the four BrakeControl Valves (BCV) to modulate required pressure to the wheelbrakes. BCVs 1 and 4 control the hydraulic pressure from system 1 tothe outboard wheels. BCVs 2 and 3 control the hydraulic pressure fromsystem 2 to the inboard wheels.

Pressure and wheel speed transducers send signals to the BCU sothat it can monitor brake performance and send the appropriate signalsto the crew alerting system and other systems. The BCU also receivessignals from the landing gear position and condition, air/groundsituation, and hydraulic system status. The system displays messageson the EICAS to indicate a failure in one pair of brakes or a failure in asingle wheel brake (brake degraded performance). In the event ofbrake system failure, the BCU will shut down the affected hydraulicsystem through the shutoff valves. The shutoff valves are energizedwhenever the landing gear is extended and de-energized after landinggear retraction.

Protective functions controlled by the normal braking system includeanti-skid protection, locked wheel protection, and touch-downprotection.

AIRPLANEOPERATIONS

MANUAL



3 Code

01

BRAKE SYSTEM SCHEMATIC


AIRPLANEOPERATIONS

MANUAL

2-12-10 Page

4 Code


ANTI-SKID PROTECTION

The anti-skid protection controls the amount of hydraulic pressureapplied by the pilots on the brakes. The anti-skid provides themaximum allowable braking effort for the runaway surface in use. Itminimizes tire wear, optimizes braking distance, and prevents skidding.

To perform this function, the BCU computes the wheel speed signalsfrom the four speed transducers. If one signals falls below the wheelspeed average, a skid is probably occurring, and braking pressure isrelieved on that side. After that wheel speed has returned to theaverage speed, normal braking operation is restored.

The anti-skid does not apply pressure on the brakes, but only relievesit. So, to perform a differential braking technique, the pilot shouldreduce pressure on the side opposite to the turn, instead of applyingpressure to the desired side.

The anti-skid system incorporates the locked wheel protection andtouchdown protection features.

LOCKED WHEEL PROTECTION

Locked wheel protection is activated for wheel speeds above 30 kt. Itcompares wheel speeds signals. If one wheel speed is 30% lower thanthat of another, a full brake pressure relief is commanded to theassociated wheel, allowing wheel speed recovery. The 30% tolerancebetween the wheel speeds is provided to permit an amount ofdifferential braking, for steering purposes.

For wheel speeds below 30 kt, the locked wheel protection isdeactivated and the brake system actuates without the wheel speedcomparator. For wheel speeds below 10 kt, the anti-skid protection isdeactivated, allowing the pilot to lock and pivot on a wheel.

AIRPLANEOPERATIONS

MANUAL



5 Code

01

DIFFERENTIAL BRAKING TECHNIQUE


AIRPLANEOPERATIONS

MANUAL

2-12-10 Page

6 Code


TOUCHDOWN PROTECTION

The touchdown protection system inhibits brake actuation before themain wheels spin up during landing. Brake actuation will be allowedonly after 3 seconds from the latest touchdown or after the wheelshave spun-up to 50 kt. In bouncing landings, the countdown is resetafter each runway contact.

Touchdown protection is provided by the brake system receivingsignals from main landing gear weight-on-wheel proximity switches. Ifone landing gear proximity switch fails at the air position, the brakesystem will operate normally. However, if both proximity switches fail atthe air position, braking capacity will be available only for wheel speedsabove 10 kt.Below 10 kt, a loss of the main brake capacity will occur, butemergency braking is still available.

GEAR-RETRACTING-IN-FLIGHT BRAKING

Gear-retracting-in-flight braking prevents the landing gear from beingretracted when the wheels are turning. This system computes signalsfrom the air/ground indicating system and from the landing gear leverposition. As soon as the airplane is airborne and the gears arecommanded to retract, it applies braking pressure to the main wheels.The nose wheels are braked by a stop within the nose landing gearwheel well.

AIRPLANEOPERATIONS

MANUAL



7 Code

01



AIRPLANEOPERATIONS

MANUAL

2-12-10 Page

8 Code


EMERGENCY/PARKING BRAKE SYSTEM

The emergency/parking brake system is used when parking theairplane or when the normal braking system has failed. Theemergency/parking brake system is mechanically commanded andhydraulically actuated. It is totally independent of the BCU, so it hasnone of the normal braking system protections.

The emergency/parking brake is controlled through a handle locatedon the left side of the control pedestal. This modulates theEmergency/Parking Brake Valve. When the Emergency/Parking BrakeValve is actuated, hydraulic pressure coming from a dedicatedaccumulator is equally applied to the four main landing gear brakes.Braking capacity is proportional to the handle displacement. A BRAKEON indicating light illuminates to indicate that pressure is being appliedto the wheel brakes. A locking device allows the handle to be held inthe actuated position, for parking purposes.

The accumulator is supplied by hydraulic system 2. A caution messageis displayed on the EICAS in case of accumulator hydraulic lowpressure. After the message is displayed, if no leakage exists, at leastone full emergency/parking brake application is available. Ifoverpressure occurs due to overheating, a thermal relief valve allowshydraulic system communication with the return. A refilling connectionis provided to allow pressurization of the accumulator.

The accumulator allows 6 complete emergency actuation or at least 24hours of parking brake actuation.

NOTE: To prevent transfer of hydraulic fluid from one system to theother, normal braking should be applied and held while theparking brake is fully applied or released.

AIRPLANEOPERATIONS

MANUAL



9 Code

01

EMERGENCY/PARKING BRAKE SYSTEM SCHEMATIC


AIRPLANEOPERATIONS

MANUAL

2-12-10 Page

10 Code

01 REVISION 7

EICAS MESSAGES

TYPE MESSAGE MEANINGEMRG BRK LO PRES Emergency/parking brake

accumulator presents alow pressure condition.

CAUTION

BRK OUTBD (INBD) INOP Outboard and/or inboardpair of brakes isinoperative.

BRAKE OVERHEAT Any brake temperaturehas exceeded 420°C.(*)

BRAKE DEGRADED Total or partial loss ofbraking capability of oneoutboard wheel (1 or 4)and/or one inboard wheel(2 or 3), or internal BCUfailure.

NOTE: (*) For EMB-135BJ airplanes equipped with LR brakes, thebrake overheat set point is 450°C.


MAIN PANEL/RAMP PANEL

1 - BRAKE ON LIGHT− Illuminates when emergency/parking brake is applied.

BRAKE ON LIGHT

AIRPLANEOPERATIONS

MANUAL



11 Code

01

CONTROL PEDESTAL

1 - EMERGENCY/PARKING BRAKE HANDLE− Actuates the emergency/parking brake valve.− Pull the handle and rotate to lock in the fully-actuated position.

EMERGENCY/PARKING BRAKE HANDLE


AIRPLANEOPERATIONS

MANUAL

2-12-10 Page

12 Code

01 REVISION 7

MFD INDICATIONS

1 - BRAKE TEMPERATURE INDICATION− Temperature is indicated by two vertical bars (one for each main

landing gear) and four pointers (one for each brake).− The scale ranges from 0 to 500°C.− The scale and pointer are green when temperature is below

200°C, and amber when equal or greater than 200°C.− The temperature indication pointer is removed from the display

in case of loss of temperature sensor signal.

NOTE: For EMB-135BJ airplanes equipped with LR brakes, the scaleand pointer are green when temperature is below 250°C, andamber when equal or greater than 250°C.

BRAKE TEMPERATURE INDICATION

AIRPLANEOPERATIONS

MANUAL



1 Code

01

NOSE WHEEL STEERING SYSTEMThe nose wheel steering system is electronically controlled andhydraulically operated. It is powered by the hydraulic system 1. TheElectronic Control Module is energized when the landing gear is downand locked, with the airplane on ground. In this condition, steering canbe controlled by either the pedals or the steering handle. In either case,the commanded displacement is measured by a potentiometer box,which transmits the signal to the Electronic Control Module. TheElectronic Control Module signals the hydraulic manifold to pressurizethe steering actuator in the commanded direction. For monitoringpurpose, a feedback potentiometer in the nose landing gear legtransmits nose wheel displacement information to the ElectronicControl Module.

Maximum nose wheel displacement values due to actuation of thesteering handle and pedals are presented in the table below in degrees:

CERTIFICATIONPEDALS

ONLY

STEERINGHANDLE

ONLY

HANDLEAND

PEDALS

CTA/JAA/FAA 5° 71° 76°

NOTE: Steering handle actuation with nose wheels beyond theiroperational limits may cause damage to the nose wheelsteering system.Check if the nose wheel position indication mark is within thenose wheel position indication scale limits.

A position sensor set to 7° disengages the system if the nose wheel isrotated above this limit by using the rudder pedals. To reengage thesystem, resume command through the handle.


AIRPLANEOPERATIONS

MANUAL

2-12-15 Page

2 Code

01 REVISION 7

The steering system may be manually disengaged through switcheslocated on the pilots' control wheels. Automatic system disablementoccurs as soon as the airplane is airborne. Nose wheel centering withthe nose gear shock absorber extension is provided by a cam. Thenose wheel is also centered by caster effect whenever the system isdisengaged.

If the Electronic Control Module detects a failure, the EICAS is signaledto present a caution message. In these cases, for airplanes Post-Mod.SB 145LEG-32-0020 or with an equivalent modification factoryincorporated, the tiller commands will be inhibited if ground speed isabove 25 kt.

Optionally, some airplanes are equipped with an external SteeringDisengagement Switch which allows ground personnel to disengagesteering prior to towing operations. The switch actuates directly on thesteering system, shutting its power down. The disengagement switchinhibits the steering actuation commanded by the steering handle andthe rudder pedals. A caution message is displayed on the EICASwhenever the steering system is disengaged by the external switch.Steering Disengagement Switch is installed in an compartment on theleft front fuselage.

EICAS MESSAGES


CAUTION STEER INOPSteering system is inoperative.Message is presented only onground.

AIRPLANEOPERATIONS

MANUAL



3 Code

01

NOSE WHEEL STEERING SCHEMATIC


AIRPLANEOPERATIONS

MANUAL

2-12-15 Page

4 Code



STEERING DISENGAGEMENT SWITH (guarded)

ENGAGED - Allows normal steering system operation.DISENGAGED - Disables steering system operation.

STEERING DISENGAGEMENT SWITCH COMPARTMENT

145AOM2120017.MCE

AIRPLANEOPERATIONS

MANUAL



5 Code

01

PILOT'S CONSOLE

1 - STEERING HANDLE− Commands nose wheel steering, allowing 71° deflection to

either side.− Push the handle down (step 1) to enable the command or to

reset the steering system after disconnection. Then rotate left orright (step 2) to command steering.

STEERING HANDLE


AIRPLANEOPERATIONS

MANUAL

2-12-15 Page

6 Code


CONTROL WHEEL

1 - STEERING DISENGAGE BUTTON− When pressed disengages the nose wheel steering system.

STEERING DISENGAGE BUTTON

AIRPLANEOPERATIONS

MANUAL



7 Code

01

EMB-135 BJ MINIMUM TURNING RADII


AIRPLANEOPERATIONS

MANUAL

2-12-15 Page

8 Code



AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-00 Page

1 Code

01

SECTION 2-06

LIGHTING

TABLE OF CONTENTS

Block Page

General .............................................................................. 2-06-05 ..01Cockpit Lighting.................................................................. 2-06-05 ..02

Controls and Indicators................................................... 2-06-05 ..04Passenger Cabin Lighting .................................................. 2-06-10 ..01

Sterile Light (Optional).................................................... 2-06-10 ..03Courtesy and Stairs Lighting .......................................... 2-06-10 ..03Controls and Indicators................................................... 2-06-10 ..04

External Lighting ................................................................ 2-06-15 ..01Service Compartments Lighting ..................................... 2-06-15 ..05Baggage Compartment Lighting..................................... 2-06-15 ..05Controls and Indicators................................................... 2-06-15 ..06

Emergency Lighting ........................................................... 2-06-20 ..01EICAS Messages ........................................................... 2-06-20 ..04Controls and Indicators................................................... 2-06-20 ..04Galley Emergency Lights Control Panel ......................... 2-06-20 ..06

LIGHTING AIRPLANEOPERATIONS

MANUAL

2-06-00 Page

2 Code



AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-05 Page

1 Code

01

GENERALThis airplane is equipped with a lighting system in order to illuminate allessential parts located inside and outside of the fuselage and to assurea proper and safe operation of the airplane.

The cockpit is illuminated by dome, chart, fluorescent/flood andreading lights.

The passenger cabin lighting is provided by the upward and downwardfluorescent lights installed in the left and right valance panels,passenger reading lights, lavatory lights and galley lights.

External lighting consists of navigation, anticollision (strobe and redbeacon), landing, taxi, inspection and logotype lights.

Emergency lights are provided inside and outside the airplane toassure, for the crewmembers and passengers, a safe nightevacuation, under emergency condition.

The system also provides lighting for baggage and servicecompartments.


MANUAL

2-06-05 Page

2 Code

01 JUNE 20, 2002

COCKPIT LIGHTINGThe lighting system inside the cockpit is composed of five differenttypes of lights, which are as follows:

- Dome lights.- Reading lights.- Chart lights.- Fluorescent flood/storm light.- Instruments and panels lights.

DOME LIGHTSCockpit illumination is provided by two dome lights of fixed intensity,commanded by a switch on the overhead panel. One light is locatedabove the pilot’s seat and the other is located above the copilot’s seat.

READING LIGHTSIn order to provide adequate light distribution for the reading of maps,check lists and manuals there are three reading lights inside thecockpit, one for the pilot, a second for the copilot and a third for theobserver.By rotating the inner bezel of each of these three light installations,lighting intensity can be adjusted from off to full bright according tocrew preference. The aperture or size of the light pattern isindependently adjustable from a small to a large square pattern byrotating the outer bezel.

CHART LIGHTSChart lights are provided to illuminate the chart holders located at thepilot’s and copilot’s control wheel.The chart light is turned on when the chart holder assembly is lifted.Light intensity is controlled by potentiometer knobs located on eachside of the glareshield panel and may be selected from dim to fullbright.

FLUORESCENT FLOOD/STORM LIGHT (OPTIONAL)Three flood/storm lights provide a proper lighting level in the cockpitand assures instrument readability when the ambient lighting is toointense with lightning flashes.The lights are located under the glareshield panel, two for the pilot’sand central side and the other for the copilot’s side. Light intensity iscontrolled by potentiometer knobs located on each side of theglareshield panel and may be selected from off to full bright.

AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-05 Page

3 Code

01

INSTRUMENTS AND PANELS LIGHTSThe instrument and control panel lights system provides lighting forinstruments, control panels, and pushbuttons. Light intensity iscontrolled by potentiometer knobs located on each side of theglareshield panel and on the overhead panel.

COCKPIT LIGHTING


MANUAL

2-06-05 Page

4 Code

01 JUNE 20, 2002


GLARESHIELD PANEL

1 - FLOODLIGHT CONTROL KNOBS− Turn on/off and regulate the brightness of flood lighting.− Pilot’s knob controls pilot’s panel, center panel and control

pedestal.− Copilot’s Knob controls copilot’s panel.

2 - CHART HOLDER LIGHTING CONTROL KNOBS− Regulate the brightness of associated chart holder lighting.

NOTE: Chart light is turned on when the chart holder assembly islifted.

3 - DISPLAYS LIGHTING CONTROL KNOBS− Regulate the brightness of Electronic Display.− Pilot’s knobs control pilot’s PFD and MFD.− Copilot’s knobs control EICAS and copilot’s PFD and MFD.

4 - PANEL LIGHTING CONTROL KNOBS− Turn on/off and regulate the brightness of panels lighting.− Pilot’s knobs control pilot’s panel, center panel and control

pedestal.− Copilot’s knob controls copilot’s panel and observer panel.

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MANUAL

2-06-05 Page

6 Code

01 JUNE 20, 2002

OVERHEAD PANEL

1 - PUSHBUTTON LIGHTS TEST SWITCH (if installed)− When actuated to the TEST position (momentary position)

allows checking of the striped bars and caption indications.− The striped bars and caption indications in all pushbuttons

located on the main panel, overhead panel, control pedestal andright lateral console will illuminate, allowing verification of lampsintegrity.

− The fire handles, APU fire extinguish button, BAGG EXTGbutton, electromechanical GUST LOCK indication lights, GPUAVAIL annunciator and digital pressurization control button willnot illuminate and will not be tested.

2 - OVERHEAD PANEL LIGHTING CONTROL KNOB− Turns on/off and regulates the brightness of the overhead panel

lighting.

3 - COCKPIT DOME LIGHTS SWITCH− Turns on/off the two cockpit dome lights.

AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-05 Page

7 Code

01

OVERHEAD PANEL


MANUAL

2-06-05 Page

8 Code

01 JUNE 20, 2002

FLIGHT CREW READING LIGHTS

1 - INNER RING− Provides turn on/off and dimming control.

2 - OUTER RING− Provides reading area adjustment, allowing light beam

orientation up to 35 degrees from the vertical axis in anydirection.

FLIGHT CREW READING LIGHTS

AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-10 Page

1 Code

01

PASSENGER CABIN LIGHTINGPassenger cabin lighting includes general illumination, reading andtable lights, lavatory, cabin signs and galley lights.

GENERAL PASSENGER CABIN ILLUMINATIONGeneral passenger cabin illumination is provided by upwash anddownwash fluorescent lights mounted in the left and right valancepanels. The forward and aft cabins can be switched independentlyfrom both VIP Control Panels and from the Galley Control Panel.Pressing one of the switches alternates the lights between Brigth, Dimand Off.

READING AND TABLE LIGHTSThe cabin reading and table lights are located in the escutcheonsabove the seats and tables. These lights are controlled by the switchesinstalled on the control panels beside each passenger seat, on thedivan surround and on the VIP Control Panels. The direction of thelight beam is fully adjustable.

LAVATORYThe Lavatory lighting consists of upwash, downwash, vanity, readingand call lights. The lavatory upwash and downwash lights installed inthe valance panels are automatically controlled through a microswitchinstalled in the latch assembly of the door. When the door is notlatched (lavatory not in use), the valance lights illuminate in reducedbrightness mode, switching to full brightness mode when the door isclosed and latched. The vanity light consists of a fluorescent tubeinstalled over the vanity mirror. A switch located in the vanity consoleturns the vanity light on or off. A switch on the lavatory control panelcontrols the reading light, located above the toilet unit. The readinglight beam is fully adjustable.


MANUAL

2-06-10 Page

2 Code

01 JUNE 20, 2002

PASSENGER CABIN SIGNSThe passenger warning signs are illuminated signs that will be clearlyvisible under normal daylight lighting conditions. They providepassengers with NO SMOKING, FASTEN SEAT BELTS, RETURN TOSEAT and DO NOT OPEN DOOR instructions. An aural signal soundswhenever any passenger sign is turned on or off by the pilot.

The NO SMOKING and FASTEN SEAT BELTS signs are controlledthrough respective switches located on the overhead panel. Thesesigns are installed along the cabin in a way they can be seen from allthe seats, in the Galley area and in the Lavatory.

In addition, a RETURN TO SEAT sign is provided only in the Galleyarea and in the Lavatory and is activated in conjunction with theFasten Seat Belts sign.

The NO SMOKING, FASTEN SEAT BELTS and RETURN TO SEATsigns are also activated when the oxygen dispensing units are openand when the cabin pressure reaches 14500 ft above sea level.

The DO NOT OPEN DOOR sign is installed beside the lavatory-to-baggage compartment access door (inside the lavatory) to warnpassengers against opening the door, whenever smoke is detected inthe baggage compartment.

GALLEY LIGHTThe galley lights are divided into three subsystems controlled throughthree switches on the Galley Control Panel. These switches turn lightson or off in each subsystem.

AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-10 Page

3 Code

01

STERILE LIGHT (OPTIONAL)A blue sterile light, located on the cockpit/pax partition, indicates, whenlit, that entry into the cockpit is not allowed. It is commanded through aswitch located at the overhead panel.

COURTESY AND STAIRS LIGHTINGThe courtesy and stair lights provide lighting for safe boarding ofcrewmembers and passengers. The courtesy and stair lights consist ofthe main door light (entry area), service door light (galley area),stairway lights and cockpit step light as follows:

− Main door light: A light is installed on the main door ceiling panel,above the entry area of the airplane, to illuminate the stair, entryarea, aisle toward cockpit and passenger cabin.

− Stairway lights: Airplanes equipped with airstair main doors havestair lights installed in each step of the main door stair to provideadequate step illumination.

− Cockpit step light: A red light is installed in the step between thepassenger cabin and the cockpit to provides light for safe entry intothe cockpit. This light is illuminated simultaneously with the maindoor light.

These lights are controlled by a main door microswitch and by twomembrane switches of the Entrance Control Panel, located in theEntertainment Cabinet, beside the main door.


MANUAL

2-06-10 Page

4 Code

01 JUNE 20, 2002



−− CABIN LIGHTING CONTROL SWITCHES The cabin lighting switches on the Galley Panel control forward andaft upwash and downwash lighting. Pressing each of these switchesalternates the lights between Bright, Dim and Off modes.


AIRPLANEOPERATIONS

MANUAL

LIGHTING

JUNE 20, 2002 2-06-10 Page

5 Code

01

VIP CONTROL PANEL

−− CABIN LIGHTING SWITCHES Cabin lighting can be controlled from both VIP Control Panels.Pressing any of the light control switches on these panels alternatesthe upwash and downwash lights between Bright, Dim and Offmodes.

VIP CONTROL PANEL (FORWARD AND AFT)


MANUAL

2-06-10 Page

6 Code

01 JUNE 20, 2002

COURTESY LIGHTING PANEL

−− COURTESY AND STAIRS LIGHTING CONTROL SWITCHES

ENTRY LIGHTS ON/OFF SWITCH

OFF (Amber) - All courtesy and stairs lights are turned off.

ON (Green) - All courtesy and stairs lights are turned on, when themain door is open. When the main door is closed, onlythe overdoor light remains on, to illuminate the maindoor area in flight.

ENTRY LIGHTS AUTO SWITCH

ON (Green) - All courtesy and stairs lights are extinguished when themain door is closed and illuminate when the main dooris open.

NOTE: The cockpit dome lights may be commanded through the

Courtesy and Stairs Lighting Control Switch provided theairplane is deenergized and the Cockpit Dome Lights Switchis set to ON position.

COURTESY LIGHTING PANEL

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OVERHEAD PANEL

1 - FASTEN SEAT BELTS AND NO SMOKING SIGNS SWITCHES− Turns on/off the associated passenger signs.

2 - STERILE LIGHT SWITCH− Turns on/off the sterile light.

OVERHEAD PANEL


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EXTERNAL LIGHTINGThe external lights necessary to a proper and safe operation of theaircraft are:

- Landing lights.

- Taxi lights.

- Navigation lights.

- Anti-collision lights.

- Wing inspection lights.

- Logotype lights.

LANDING LIGHTS

The landing lights provide adequate lighting during final approach,flare-out and take-off. Two landing lights are fitted in the wing leadingedge close to the fuselage and a third landing light is mounted on thenose landing gear strut. The switches located on the overhead panelare responsible for the control of the landing lights.

TAXI LIGHTS

The taxi light provides sufficient intensity and beam spread to aid pilotsduring all taxi operation phases, covering the runway and adjacentareas.Two taxi lights are fitted on the nose landing gear strut and arecommanded by a single switch located on the overhead panel.

NAVIGATION LIGHTS

The navigation lights, red on the left and green on the right, are fitted tothe leading edge of each wing tip. A white navigation light is fitted tothe cone top of the horizontal stabilizer. The navigation lights arecontrolled by a switch located on the overhead panel.

ANTI-COLLISION LIGHTS

The anti-collision lights provide illumination for visual recognition andcollision avoidance during all flight/taxi operations. White strobe (anti-collision) lights are fitted to each wing tip and cone top of the horizontalstabilizer. Red beacon lights are mounted on the upper fuselage. Twodifferent switches, one for strobe lights and another for the red beaconlights are located on the overhead panel.


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WING INSPECTION LIGHTS

Two inspection lights, one on each side of the fuselage, providelighting of the wing leading edge to allow the crew to verify iceformation. The inspection lights are controlled by a switch located onthe overhead panel.

LOGOTYPE LIGHTS (OPTIONAL)

The logo lights are installed on the underside of the horizontal stabilizerand are aimed at the vertical fin. They provide adequate illumination ofthe airplane’s logo during operation on the ground and in flight. Aswitch located on the overhead panel controls the logotype lights.

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SERVICE COMPARTMENTS LIGHTING

The system provides lighting in the service compartments for quickinspection and accomplishment of several tasks. Service lights areinstalled in the nose landing gear, rear and forward electronic bays, tailcone and forward flight control compartments. The lights are controlledby a door micro-switch, that turns on the associated light when theaccess doors is open, or by dedicated switches, installed in thecompartment.

BAGGAGE COMPARTMENT LIGHTING

The baggage compartment is equipped with lights installed on theceiling panel. These lights are controlled by a microswitch installed atthe door structure and they are turned on whenever the baggagecompartment door or the lavatory-to-baggage compartment door isopen.

In addition, The airplane is equipped with a cargo door light installed inthe left pylon. The cargo door light provides lighting externally of thecargo compartment and is controlled by a microswitch when theaccess door of the cargo compartment is open.


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OVERHEAD PANEL

1 - NAVIGATION, RED BEACON, STROBE AND WINGINSPECTION LIGHTS SWITCHES− Turns on/off the associated light.

2 - LOGOTYPE LIGHTS SWITCH (OPTIONAL)− Turns on/off the logotype lights.

3 - TAXI LIGHTS SWITCH− Turns on/off the taxi lights.

NOTE: Taxi lights are not turned on if nose landing gear is notdown and locked, regardless of the Taxi Lights Switchposition.

4 - LANDING LIGHTS SWITCHES− Turn on/off the associated landing light.

NOTE: Nose landing light is not turned on if nose landing gear isnot down and locked, regardless the of Nose Landing LightSwitch position.

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EMERGENCY LIGHTINGThe emergency lighting consists of internal and external lights thatprovide proper illumination for emergency cabin evacuation. Theselights are powered by four dedicated batteries charged through theEssential Bus. Batteries power is sufficient to supply all internal andexternal emergency lights for approximately 15 minutes.

The exterior emergency lights installed are as follows:− Two lights are installed on the right hand wing to fuselage fairing, in

order to illuminate the wing escape route and the ground area.

− One light is installed on the handrail support of the main door inorder to illuminate the outside area where the evacuee is likely tomake first contact with the ground.

Internal emergency lights consist of the cockpit light, aisle lights, maindoor lights, overwing emergency exit lights, floor proximity lights andEXIT signs as follows:

− Cockpit light: This light is located on the cockpit ceiling to providegeneral cockpit emergency illumination.

− Aisle lights: Four dome lights are located along the aisle for generalemergency cabin illumination.

− Main door and overwing emergency exit lights: One light is installedin the ceiling panel over the main exit door illuminating the main exitdoor area and the outside area. An emergency light is installedbehind the valance panel, above the overwing emergency exit,providing illumination over the exit handle instructions.

− Floor proximity emergency lights: Electro-luminescent stripindicators are installed on the inboard side of the seats, in the toekick area of the credenza and cabinets to provide a means ofidentifying the emergency escape path even in conditions of densesmoke.

− Illuminated EXIT signs: One emergency Exit sign is installed nearthe main door and two others are installed near the overwing exitdoor. In addition, Exit locator signs are located on the forward righthand bulkhead and aft left hand bulkhead of the forward cabin. Theaft cabin exit locator sign is located on the forward left handbulkhead.


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Emergency lighting is controlled through the Emergency Lights Switch,located on the overhead panel, and through the equivalent membraneswitch located on the Galley Emergency Lights Control Panel.

A caution message is presented on the EICAS if the system is notarmed.

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AREA ILLUMINATED BY EMERGENCY LIGHTING


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EICAS MESSAGE

TYPE MESSAGE MEANINGCAUTION EMERG LT NOT ARMD Emergency lighting system is

not armed.


OVERHEAD PANEL

1 - EMERGENCY LIGHTING SWITCHON - Emergency lights illuminate with power supplied by the

dedicated batteries.ARM - Emergency lights are in standby mode (lights turned off and

the batteries being charged) and illuminate automatically incase of an electrical emergency, with power supplied by thededicated batteries.

OFF - Emergency lights are turned off. Emergency lightingdedicated batteries are not charged.

NOTE: The emergency lights are controlled by the EmergencyLights Switch on the overhead panel only when the color ofthe Emergency Lights Switch located on the Galley cabinet(normal flight position).

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GALLEY EMERGENCY LIGHTS CONTROL PANEL

AMBER - Emergency lights remain in the mode selected by theEmergency Lighting Selector Knob position on the cockpit(normal flight position).

GREEN - Emergency lights are turned on with power supplied bydedicated batteries, regardless of the Emergency LightingSelector Knob position on the cockpit.

GALLEY EMERGENCY LIGHTS CONTROL PANEL

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SECTION 2-10

POWERPLANTTABLE OF CONTENTS

Block Page

Index ................................................................................. 2-10-00 ..01

General .............................................................................. 2-10-05 ..01Main Assemblies ............................................................ 2-10-05 ..02

Fan Module ................................................................. 2-10-05 ..02High-pressure Compressor ........................................ 2-10-05 ..02High-pressure Turbine (HPT) ..................................... 2-10-05 ..02Low-pressure Turbine (LPT)....................................... 2-10-05 ..02Exhaust Cone and Mixer ............................................ 2-10-05 ..02Accessory Gearbox .................................................... 2-10-05 ..03

Engine Fuel System ........................................................... 2-10-10 ..01Fuel Pump and Metering Unit (FPMU) ........................... 2-10-10 ..01Fuel Cooled Oil Cooler (FCOC)...................................... 2-10-10 ..02Compressor Variable Geometry Actuation System........ 2-10-10 ..02Fuel Nozzles ................................................................... 2-10-10 ..02

Lubrication System............................................................. 2-10-15 ..01Lubricating Oil Supply System........................................ 2-10-15 ..01

Oil Tank ...................................................................... 2-10-15 ..01Lube and Scavenge Pump ......................................... 2-10-15 ..02Oil Filter Unit ............................................................... 2-10-15 ..02Air-Cooled Oil Cooler (ACOC) .................................... 2-10-15 ..02Fuel-Cooled Oil Cooler (FCOC).................................. 2-10-15 ..02

Engine Sumps ................................................................ 2-10-15 ..03Lubricating Oil Scavenge System................................... 2-10-15 ..03Lubricating Oil Vent System ........................................... 2-10-15 ..03

Engine Bleed...................................................................... 2-10-20 ..01

Engine Electrical System ................................................... 2-10-25 ..01Electrical Power Sources................................................ 2-10-25 ..01Permanent Magnet Alternator (PMA) ............................. 2-10-25 ..01

Ignition System................................................................... 2-10-30 ..01Pneumatic Starting System................................................ 2-10-30 ..02

Air Turbine Starter (ATS)................................................ 2-10-30 ..02Starting Control Valve (SCV).......................................... 2-10-30 ..02Starting By Using Ground Equipment............................. 2-10-30 ..03

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Engine Indicating System (EIS)..........................................2-10-35.. 01Engine Sensors ..............................................................2-10-35.. 01

Pressure/Temperature Transducer Sensor ................2-10-35.. 01Low Oil-Pressure Sensor ............................................2-10-35.. 01Oil-Level and Low-Level Sensor .................................2-10-35.. 01Electrical Oil-Filter Impending-Bypass Indicator .........2-10-35.. 01Fuel Temperature Sensor ...........................................2-10-35.. 02Electrical Fuel-Filter Impending-Bypass Indicator.......2-10-35.. 02Magnetic Indicating Plug .............................................2-10-35.. 02Igniter Spark-Rate Detector ........................................2-10-35.. 02Vibration Sensors........................................................2-10-35.. 02Fuel Flowmeter ...........................................................2-10-35.. 02

Powerplant Control System ................................................2-10-40.. 01Full Authority Digital Electronic Control (FADEC) ...........2-10-40.. 01N1TARGET Calculation..................................................2-10-40.. 04N1REQUEST Calculation ...............................................2-10-40.. 04Ground/Flight Idle Thrust Schedule ................................2-10-40.. 05Closed-Loop Fan Speed Control ....................................2-10-40.. 05N1/N2 Overspeed/Underspeed Protection .....................2-10-40.. 06Interstage-Turbine Temperature (ITT) Limiting ..............2-10-40.. 06Acceleration/Deceleration Limiting .................................2-10-40.. 06Flameout Detection/Autorelight ......................................2-10-40.. 06N1 Reversionary Control Mode.......................................2-10-40.. 07FADEC Inputs Selection and Fault Accommodation ......2-10-40.. 07FADEC Discrete Outputs................................................2-10-40.. 07Alternate FADEC Selection.............................................2-10-40.. 08FADEC Reset .................................................................2-10-40.. 08

Engine Operation................................................................2-10-50.. 01General ...........................................................................2-10-50.. 01Thrust Ratings ................................................................2-10-50.. 01Engine Control ................................................................2-10-50.. 02Thrust Management........................................................2-10-50.. 02

Thrust Mode Selection ................................................2-10-50.. 02AE3007A1E Thrust Mode Selection............................2-10-50.. 07Fan-Speed Scheduling................................................2-10-50.. 08Alternate Takeoff Thrust Control System (ATTCS) ....2-10-50.. 10Takeoff Data Setting ...................................................2-10-50.. 11

Engine Start ....................................................................2-10-50.. 14Engine Dry Motoring....................................................2-10-50.. 15

Engine Shutdown............................................................2-10-50.. 15EICAS Messages ...............................................................2-10-50.. 16

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Controls and Indicators ...................................................... 2-10-60 ..01Control Pedestal ............................................................. 2-10-60 ..01Powerplant Control Panel............................................... 2-10-60 ..03Fire Handle ..................................................................... 2-10-60 ..05Engine Indication on EICAS ........................................... 2-10-60 ..05Takeoff Page on MFD .................................................... 2-10-60 ..10First Engine Backup Page on RMU................................ 2-10-60 ..11

Thrust Reverser ................................................................. 2-10-70 ..01General........................................................................... 2-10-70 ..01Lock Protection............................................................... 2-10-70 ..01Operation........................................................................ 2-10-70 ..01

Operation Logic........................................................... 2-10-70 ..02EICAS Indication......................................................... 2-10-70 ..02

Thrust Reverser Interlock ............................................... 2-10-70 ..03EICAS Messages ........................................................... 2-10-70 ..03


MANUAL

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GENERALThe airplane is powered by two fuselage-mounted Allison turbofanengines. Engine denominations, thrust (installed, static sea level) andflat rates are as follows:

ENGINE MODEL MAX. T/O THRUST FLAT RATEAE3007A1E EMB-135 BJ 8810 lb ISA+19°CAE3007A1P EMB-135 BJ 8169 lb ISA+19°C

NOTE: -Max T/O thrust and flat rate values for AE3007A1P are based on T/O RSV thrust.- Max T/O thrust and flat rate values for AE3007A1E are based on E T/O RSV thrust.

The AE3007 is a high bypass ratio, two-spool axial flow turbofanengine. The main design features include:

− A single stage fan,− A 14-stage axial-flow compressor with inlet guide vanes and five

variable-geometry stator stages,− A 2-stage high pressure turbine to drive the compressor,− A 3-stage low pressure turbine to drive the fan,− Dual, redundant, Full Authority Digital Electronic Controls

(FADEC),− Accessory gearbox,− Air system for aircraft pressurization and engine starting.

Each engine is controlled by redundant FADECs. The FADECs alsoprovide information to the EICAS, although some parameters signalsare provided directly from engine sensors. All powerplant parametersare indicated on the EICAS, which also provides warning, caution andadvisory messages.

The cockpit control stand incorporates two thrust levers, one for eachengine, and four buttons for engine thrust rating selection.

Controls for ignition, FADEC, takeoff data setting, takeoff ratingselection and engine Start/Stop are located on the overhead panel.


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MAIN ASSEMBLIES

FAN MODULE

Air enters the engine through the fan case inlet and is compressed bya 24-blade, single-stage fan. The compressed air is split into a bypassstream, which bypasses the core through the outer bypass duct, and acore stream that enters the high-pressure compressor.

HIGH-PRESSURE COMPRESSOR

The compressor rotor consists of 14 stages of individual wheelassemblies, compressor shaft, compressor-to-turbine shaft, andcompressor tiebolt. Compressor Variable Geometry (CVG) stators areprovided for stages 1 through 5 and for the inlet guide vanes. Thesestators are driven by servo actuators controlled by the FADECs. High-pressure compressor bleed air tappings are available at the 9th and14th stages (compressor discharge). A combustion liner assembly mixes air and fuel to support combustion,and delivers a uniform, high-temperature gas flow to the turbine.

HIGH-PRESSURE TURBINE (HPT)

The High Pressure Turbine converts the gas flow coming from thecombustion liner into usable mechanical energy to drive thecompressor.

LOW-PRESSURE TURBINE (LPT)

The Low-Pressure Turbine is located downstream of the High-Pressure Turbine and extracts energy from the gas path to drive thefan. The LPT is connected to the fan by means of a shaft extendingthrough the entire high-pressure spool and the compressor assembly.Air exiting the LPT mixes with the bypass air and provides thrust.

EXHAUST CONE AND MIXER

The mixer provides the mixing chamber for the engine bypass andcore gas-flow streams and sets the fan operating line for all operatingenvelope conditions. The Thrust Reversers deflect the exhaustproviding reverse thrust.

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ACCESSORY GEARBOX

An accessory gearbox is driven by the high-pressure spool andprovides driving pads for the following engine and airplaneaccessories:

− Engine accessories: Fuel Pump and Metering Unit (FPMU),Permanent Magnet Alternator (PMA), and oil pump.

− Airplane accessories: hydraulic pump, electrical generators, andpneumatic starter.

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ENGINE FUEL SYSTEMThe Engine Fuel System has a distribution and an indicating system.The distribution system supplies filtered and metered fuel forcombustion. Secondary functions include providing pressurized fuel toactivate the Compressor Variable Geometry (CVG) system, andproviding a cooling medium for lubrication oil. The indicating systemcomponents monitor the fuel supply and are located on the engines.

The engine fuel system comprises a Fuel Pump and Metering Unit(FPMU), a Fuel Cooled Oil Cooler (FCOC), a Compressor VariableGeometry (CVG) actuator and fuel nozzles.

FUEL PUMP AND METERING UNIT (FPMU)

The FPMU is an electrical-mechanical, fully-integrated line replaceableunit which incorporates the engine fuel pumping, filtering, and meteringfunctions, and operates under authority of the engine FADECs. TheFPMU controls and supplies fuel to the engine nozzles at correctpressure and flow rate for engine start, correct engine operation,engine stop, and also controls the compressor variable-geometryvanes.

The pump system contains a low-pressure centrifugal pump and ahigh-pressure gear pump. The centrifugal pump raises the pressure ofincoming fuel high enough to meet the inlet pressure requirements ofthe high-pressure pump, with allowances for pressure losses in the fuelfilter and the FCOC. The centrifugal pump also provides vapor-freefuel to the gear pump.

The main fuel filter, located upstream of the gear pump, protects thepump metering unit components and fuel nozzles from fuelcontaminants. A fuel flow bypass valve allows continued operation inthe event of complete filter blockage.

A fuel flow pressure relief valve across the pump protects the fuelsystem from overpressure conditions.

An air vent valve provides automatic venting of entrapped air or fuelvapor at the gear pump discharge during engine starting and/ormotoring. The vent valve remains closed whenever the vent solenoid isnot energized, thus preventing fuel leakage through the vent system ifthe airplane boost pumps are turned on while the engine is not running.


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The fuel-metering valve is controlled by the FADEC and controls fueldistribution from the gear pump to the engine fuel nozzles.

Downstream of the metering valve, a pressurizing valve (PRV)generates adequate system pressure for the proper functioning of themain metering valve and pressure drop servos and CVG hydraulicactuator. The PRV also provides the primary means for engine fuelshutoff, commanded through the Latching Shutoff Valve, that receivesa Stop cockpit input through the FADEC.

FUEL-COOLED OIL COOLER (FCOC)

The FCOC is installed externally on the botton of the outer bypassduct, aft region. Fuel flows from the FPMU’s centrifugal pump to theFCOC where it simultaneously cools the engine's lubrication oil andwarms the fuel. A thermal/pressure bypass valve bypasses oil flow toprevent fuel leaving the FCOC from being heated above 93.3°C(200°F). The oil is also bypassed if the differential oil pressure isgreater than 50 psi due to hung or cold starts. After the FCOC, the fuelgoes to the filter.

COMPRESSOR VARIABLE GEOMETRY (CVG)ACTUATION SYSTEM

The high-pressure compressor has a variable geometry vane systemon its five stages to provide maximum engine performance over a widerange of engine speeds. The FADEC contains a schedule of vanepositions versus corrected gas generator speed (N2) that has beenselected to provide the optimum compressor efficiency of steady-stateconditions and adequate stall margins during transients.

The FADEC senses the vane position and, by means of fuel pressurefrom the FPMU, commands the CVG actuator movement to positionthe compressor-inlet guide vanes and the first five rows of compressorvanes to the desired setting.

FUEL NOZZLES

Each engine has 16 fuel nozzles, that furnish atomized fuel to thecombustor at the proper spray angle and pattern, for varying airflowconditions.

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LUBRICATION SYSTEMThe engine lubrication system is a self-contained, pressure-regulatedand recirculating dry sump system. The system supplies filtered andpressurized oil to the various engine oil coolers, engine sumps and theaccessories gearbox, at the proper temperature, to cool and lubricatethe bearings, seals, and gear meshes.

The main subsystems of the oil system are: lubricating oil-supply,engine sumps, lubricating oil scavenge and lubricating oil vent.

LUBRICATING OIL-SUPPLY SYSTEM

Oil is supplied to the lube and scavenge pump from a pressurized oiltank and is pumped through an oil filter. The oil is then cooled whilepassing through two heat exchangers (ACOS and FCOC). Oil pressureis controlled by a pressure-regulating valve in the pump housing. Atank pressurizing valve maintains positive pressure in the oil tank toensure an adequate oil supply to the lube and scavenge pump, andproper oil pressure at altitude. A separate Tank Vent Valve protects thetank from over-pressurization. Oil to the accessory gearbox isdistributed through cast passages to the various gear meshes andbearings. Pressurized oil is divided inside the front frame and routed tothe fan and front sumps. An external tube delivers oil from the frontframe to the compressor diffuser and the rear turbine-bearing support.

The main components of this subsystem are as follows: oil tank, lubeand scavenge pump, oil filter unit, air-cooled oil cooler (ACOC) andfuel-cooled oil cooler (FCOC).

OIL TANK

The oil tank is designed to store a sufficient amount of oil (12 quarts)for lubrication of the engine and the accessory gearbox. The tank hasan oil level sight gage and an oil level/low level warning sensor. Thesesensors allow the oil level to be continuously read remotely, andincludes a switch that is actuated when there are 6 quarts or less ofusable oil remaining in the tank. A screen on the oil outlet and a chipcollector plug at the tank bottom are protective devices that preventdebris from recirculating. The tank is protected from over-pressurization by the externally vented Pressure Relief Valve.


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LUBE AND SCAVENGE PUMP

The pressure and scavenge pumps are all mounted in a single integralunit. A single shaft drives six pumping elements. One pressurepumping element pumps oil from the tank to the system and fivescavenge pumping elements pump oil from the sumps and thegearbox to the oil tank. The pump assembly also includes a pressureregulating valve which controls oil pressure. Main Oil Pressures varieswith the center sump air pressure. A line connecting one side of theregulating valve to the center sump enables the regulating valve tocompensate for the air pressure inside the sump.

OIL FILTER UNIT

The filter unit includes a replaceable filter element, and mechanicaland electrical impending-bypass indicators. A bypass valve opens andallows oil to bypass the filter during cold starts, or when the filterbecomes excessively contaminated. A screen is located in the bypassinlet to prevent passage of particles. The electrical impending-bypassindicators provide the remote monitoring of the system.

AIR-COOLED OIL COOLER (ACOC)

The ACOC is a surface-type heat exchanger with a single plate-fin oilsection. Filtered, pressurized oil enters a manifold and flows throughthe air-cooled heat exchanger. A thermal/pressure bypass valvesenses ACOC outlet temperature. When open, this valve allows coldoil to bypass the ACOC and, once closed, forces hot oil to flow throughthe cooler. The bypass valve also opens if the cooler is obstructed.

FUEL-COOLED OIL COOLER (FCOC)

The FCOC is a heat exchanger that simultaneously cools the enginelubrication oil and warms the fuel upstream of the FPMU filter. Athermal/pressure bypass valve prevents fuel overheat. This valve alsoopens in case of cooler obstruction or cold starts.

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ENGINE SUMPS

There are four engine sumps that encompass five main-shaft bearings,four bevel-gear bearings, and six carbon seals. These sumps are asfollows: fan sump, front sump, center sump and aft sump.

LUBRICATING OIL SCAVENGE SYSTEM

Air and oil are removed from each of the sumps and directed toindividual scavenge inlets on the oil pump. The scavenge section ofthe pump includes five pumping elements and has separate inlets foreach of the engine sumps and the accessory gearbox. Each of thesump inlets to the pump includes a debris monitor with magnetic chipcollector and screen in order to protect the pumping elements. Thegearbox sump inlet to the pump contains only a screen.

LUBRICATING OIL VENT SYSTEM

All the engine sumps are vented to the accessory gearbox. The oil tankalso vents to the gearbox through a core-external line that contains atank-pressurizing valve. A Tank Vent Valve is located upstream of thepressurizing valve and is vented to the atmosphere.The gearbox acts as an air/oil separator removing any oil contained inthe vent air. The air vented by the gearbox breather is conductedthrough a transfer tube and dumped to the core exhaust.

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ENGINE BLEEDAir is bled from the compressor 9th stage during engine starting toassist with accelerating to idle rpm.

There are two different types of compressor acceleration bleed valves(CABV). The original type used two valves per engine locatedexternally on the HP compressor at approximately the 12:00 and 6:00O´clock positions. The second type is a single valve at 6:00 O´clockposition.

The engine also provides bleed air to the Pressurization and AirConditioning system through the Engine Bleed Valve (EBV), Bleed airfor this system is extracted from the 9th or 14th stages depending onthe request. Refer to section 2-14-05 for more information.


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ENGINE ELECTRICAL SYSTEM

ELECTRICAL POWER SOURCES

Primary electrical power for engine control and the ignition system isprovided by a permanent magnet alternator (PMA) that is driven by theengine accessory gearbox. Before the PMA attains sufficient speed togenerate electrical power, the airplane 28 V DC system is used topower the FADEC. Aircraft 28 V DC is also used to energize a fail-safeignition relay, so that in the event of aircraft power loss the ignition isturned on and the air vent valve is closed, thus preventing fuel leakagethrough the vent port.The PMA is the only source of power for the igniters. If a PMA failureoccurs there will not be any spark from the igniters.

PERMANENT MAGNET ALTERNATOR (PMA)

The PMA provides electrical power for both engine FADECs and to theredundant ignition systems.The PMA provides sufficient power to drive the ignition system at allspeeds above 10% N2, and powers the FADECs at a minimum of50% N2. The PMA also provides power to the Thrust Rating ModeButtons, in case of electrical emergency.For starting and emergency backup, the engine control systemrequires aircraft supplied 28 V DC (GPU and/or batteries) power.


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IGNITION SYSTEMThe engine has a dual redundant ignition system composed of twoignition exciters, two high-tension igniter leads and two igniters.

The ignition system is turned on by the FADEC during engine startingcycle or when an engine flameout condition is detected (auto-relight).

Each ignition exciter is controlled by a separate FADEC and poweredby a separate electrical winding of the PMA.

Continuous ignition or ignition off can be manually selected through theIgnition Selector Knob, located on the Powerplant Control Panel andconnected to the FADECs. Ignition control is performed according toIgnition Selector Knob position, as follows:

− Ignition Selector Knob set to ON:− Both FADECs command associated ignition channel during

start, as soon as the PMA provides sufficient power.− The ignition is not automatically deactivated when the start

cycle is completed.− If the engine is already running, both FADECs activate their

ignition channels.

− Ignition Selector Knob set to AUTO:− During ground start, only the FADEC in control activates the

ignition system at the proper time. The engine start will beperformed with only one exciter. The exciters will bealternately selected for each subsequent ground start.

− The FADEC deactivates the ignition system when the enginestarting cycle is completed.

− The auto-relight function activates the ignition system.

− Ignition Selector Knob set to OFF:− If the engine is not running, the FADEC neither activates the

ignition system nor actuates the engine fuel valve fromclosed to open position.

− If the engine is already running, at least in IDLE thrust, theFADEC does not close the engine fuel valve.


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PNEUMATIC STARTING SYSTEMThe engine starting system comprises the Air Turbine Starter and theStarting Control Valve. The starting system has the function ofsupplying airflow for pneumatic engine starting, converting thepneumatic energy into gearbox driving torque.

Pneumatic power source can be selected from the APU, ground airsupply source, or cross bleed from the opposite engine.

AIR TURBINE STARTER (ATS)

The ATS is installed in a dedicated engine accessory gearbox pad andconsists basically of an air inlet, an impeller turbine, a reductiongearset, a clutch, and an output shaft.

The ATS converts pneumatic energy into driving torque for engine gasgenerator spool acceleration up to the self-sustained speed during thestarting cycle. The air exhaust from the turbine is discharged into theengine nacelle compartment.

STARTING CONTROL VALVE (SCV)

The SCV regulates the pressure supplied to the ATS and providesisolation from the pneumatic system following start completion. Thevalve is electrically controlled and pneumatically actuated.

A SCV visual position indication is available on the valve housing.

A manual override adapter is available on the valve housing, enablingengine start in the case of a valve or associated electrical systemfailure. The valve is spring-loaded to the closed position.

If the ATS shutoff valve remains open after 53% N2, a cautionmessage is presented on the EICAS.

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STARTING BY USING GROUND EQUIPMENT

The system is pressurized by a pneumatic ground equipmentconnected to start the engine 2.

The SCV energizes to open when a starting switch ground signalenergizes the engine 2 start relay.

When the engine gas generator attains 53% N2, a validation time of 10seconds elapses before the message “E2 ATS SOV OPN” appears onthe EICAS. At 54.6% N2 the FADEC sends a signal to engine 2 startrelay be de-energized, thus the SCV is also de-energized and theairflow stops flowing to the ATS turbine. In normal operationconditions, 54.6% N2 is reached in less than 10 seconds.

The ATS stops operating and the engine gas generator speedincreases.


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ENGINE INDICATING SYSTEM (EIS)

The EIS is composed of a wiring harness and a set of engine-mountedsensors. This system is directly connected to the EICAS, providing realtime monitoring of the engine oil, fuel, and mechanical systems.

ENGINE SENSORS

PRESSURE/TEMPERATURE TRANSDUCER SENSOR

This sensor combines engine oil pressure and temperaturetransducers in a single housing, mounted on the Fuel-Cooled OilCooler (FCOC). The pressure and temperature transducers areelectrically independent and require separate signal conditioning.Due to the characteristic of some pressure sensors, the EICAS maydisplay approximately 90 psi for a 2 minutes period, for actualpressures between 90.5 and 155 psi. Considering this characteristic,pressure indication may jump suddenly from approximately 90 psi tothe actual pressure value, after the 2 minutes period is expired.

LOW OIL-PRESSURE SENSOR

The function of the low oil-pressure sensor is to give an indicationwhen oil pressure is low. This sensor is also mounted on the FCOC. Awarning message is presented on the EICAS in case of low oilpressure.

OIL-LEVEL AND LOW-LEVEL SENSOR

The engine oil-level sensor is a transducer located in the oil tank thatgives continuous and accurate oil level readings from 3qts to 12qts.The low-level warning sensor is electrically open with 6qts or less of oilremaining in the tank and remains closed otherwise. An indication ofoil-level is provided on the Takeoff page on the MFD. The indicationturns amber when oil level is at 6 quarts or below.

ELECTRICAL OIL-FILTER IMPENDING-BYPASS INDICATOR

The engine electrical oil-filter impending-bypass indicator is located inthe oil-filter assembly. An advisory message is presented on theEICAS if the differential pressure across the oil filter exceeds its setpoint.


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FUEL TEMPERATURE SENSOR

The engine fuel-temperature sensor has an indication range of -54° to176°C (-65° to 350°F) and is located on the FCOC. A caution messageis presented on the EICAS in case of fuel low temperature (below 5°Cin the engine).

ELECTRICAL FUEL-FILTER IMPENDING-BYPASS INDICATOR

The engine electrical fuel-filter impending-bypass indicator is locatedon the engine fuel pump and metering unit (FPMU). An advisorymessage is presented on the EICAS if the differential pressure acrossthe filter exceeds its set point.

MAGNETIC INDICATING PLUG

The magnetic indicating plug is located in the oil tank. The magneticplug contacts are normally open and are electrically closed whenconductive material bridges the gap between them.

IGNITER SPARK-RATE DETECTOR

The engine igniter spark-rate detectors are outputs from the ignitionexciters that indicate that an electric field has collapsed in the excitercircuit. A signal is available for each igniter circuit on the engine.

VIBRATION SENSORS

The engine vibration sensors are accelerometers that detect abnormalfan rotor and turbine rotor vibration. The transducers are connectedthrough the engine wiring harness to the EICAS.

FUEL FLOWMETER

The fuel flowmeter is a turbine, mass flow sensor. A given fuel flowthrough the sensor causes the turbine to move to a calibrated position,providing a specific voltage output to the Data Acquisition Unit (DAU).The DAU converts the voltage signal from the sensor into a flow-ratevalue (pounds or kilograms per hour) for cockpit display. The fuelflowmeter is calibrated for a range between 130 to 4300 pph. Duringsome starts, fuel flow may drop to values out of the flowmeter range. Inthis case a zero fuel flow will be displayed on EICAS for a fewseconds.

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POWERPLANT CONTROL SYSTEMEach AE 3007A engine series features a dual redundant electroniccontrol system. The main components of the powerplant controlsystem are the Full Authority Digital Electronic Controls (FADECs), theFPMU, the Permanent Magnetic Alternator (PMA), the ControlPedestal and the Powerplant Control Panel.

Thrust management logic schedules a corrected fan speed (N1) basedon a signal from the ADC and cockpit, sending it to engine controllogic, which controls the engine fuel flow and compressor variablegeometry (CVG) to attain the required engine steady-state andtransient response.

Engine control logic also incorporates engine protection logic thatprevents engine damage attributable to excessive rotor speed at alltimes, and temperature limits after the engine has completed a start.

FULL AUTHORITY DIGITAL ELECTRONIC CONTROL(FADEC)

Each engine is controlled by one of two FADECs that are namedFADEC A and FADEC B. All signals between each FADEC and itsrespective engine and between the FADECs and the airplane arecompletely redundant and isolated. This allows either A or B FADEC tocontrol the engine independently.

The FADECs are interconnected by dedicated Cross-Channel DataLinks. These buses are used to transmit engine data and FADECstatus between the two FADECs.

Each FADEC is connected to one of the two FADECs on the oppositeengine via data bus. Across this bus, the FADECs communicate theinformation necessary to implement thrust reverser interlock andAutomatic Takeoff Thrust Control System (ATTCS).

Airplane electrical power is fed to the FADEC for engine start as a solepower source until N2 is approximately 50%. Primary electrical powersource for each FADEC is generated by a dedicated set of windings inthe permanent magnet alternator (PMA). The airplane power source isfed the FADEC as a backup in the event of a failure in the PMA. In theevent of total loss of airplane power the pilot would control the enginenormally.


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Each FADEC receives command signals from the Control Pedestaland from the Powerplant Control Panel and sends a command signalto the FPMU, which meters the fuel flow to the engine in order to reachthe fan spool speed calculated by the FADEC thrust managementsection.

Both FADECs alternate powerplant control. While one FADEC controlsthe powerplant, the other remains in standby mode. The standbyFADEC monitors all inputs, performs all computations, and performsbuilt-in-test and fault detection. However, the output drivers (fuel flowand CVG control), that command the engine, are powered off.

The active FADEC is alternated at each engine ground start in order tominimize the probability of latent failure within the powerplant controlsystem/airplane interface.

The selection logic resides within the FADECs that memorize whichFADEC was used for the last engine start and commands the otherone to perform the next start, regardless of which FADEC is used inflight.

For example: If FADEC B was used for the last start, when the pilotactuates the next start, the selection logic will select FADEC A, asshown in the following table:

Start In flight (alternated) Following start

FADEC A FADEC B or A FADEC B

FADEC B FADEC A or B FADEC A

Transfer from active FADEC to standby FADEC may also beaccomplished automatically, in response to a detected fault, ormanually, through the FADEC Selector Knob, located on the overheadpanel. The manual selection overrides the automatic selection of thecontrolling FADEC unless the manually selected FADEC is not capableof safely controlling the engine.

If a fault condition is detected in the engine sensor, actuator interface,or airplane interface of the controlling FADEC, it will maintain control byusing data borrowed from the standby FADEC. If required data is notavailable, the controlling FADEC will use default data or switch toreversionary control mode.

Control will be transferred to the standby FADEC only when thecontrolling FADEC detects a fault that will result in degraded engineoperation or will render it unable to control the engine.

All measured powerplant control parameters, control system faults andstatus information are presented on the EICAS.

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N1TARGET CALCULATION

The FADEC calculates the maximum available engine thrust for agiven thrust rating mode, airspeed and ambient conditions, and bleedair configuration. Maximum thrust corresponds to N1TARGETdisplayed on the EICAS as a cyan bug on the N1 analogic indicatorarc.

When the Thrust Lever is set to the THRUST SET position, theFADEC controls the engine at N1TARGET.

In normal mode (with no ADC faults) the following data are used asprimary reference for the N1TARGET calculation:

− Pressure Altitude and Mach Number reference from ADCs.− Temperature references (REF TO TEMP during takeoff and

ADC TAT in flight).− A-ICE condition (REF A-ICE during takeoff and actual A-ICE

system feedback in flight).− Takeoff mode.

N1REQUEST CALCULATION

The N1REQUEST is a function of N1TARGET and Thrust Lever Angle.The FADEC controls the engine to N1REQUEST at steady state,except if the thrust lever is at Ground Idle position. In this case, theengine is controlled according to the Ground Idle N2 schedule.

The table below presents the main Thrust Lever positions,corresponding Thrust Lever Angle bands, and N1REQUEST forground operation.

POSITION TLA N1REQUESTMAX REVERSE 0 to 4° N1REVMIN REVERSE 14° to 22° N1IDLE

IDLE 22° to 28° N1IDLETHRUST SET 72° to 78° N1TARGETMAX THRUST Above 78° N1TARGET

N1REV is the N1 value for MAX REVERSE thrust.

Each thrust lever modulates engine thrust linearly between IDLE andTHRUST SET position. There is no thrust modulation between IDLEand MIN REVERSE.

N1REQUEST is shown as a green bug on the N1 analogic indicationarc on the EICAS.

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GROUND/FLIGHT IDLE THRUST SCHEDULE

There is only one IDLE position on the thrust lever control pedestal.However, there are two different IDLE ratings (ground and flight Idle),set as a function of the Air/Ground input to the FADEC:

− GROUND IDLE SPEEDDuring ground operations, the FADEC commands the engine toGround Idle Speed, which is scheduled in order to:− Avoid engine flameout, overtemperature or inability to accelerate.− Provide the required air bleed flow pressure and temperature for

the ECS.− Provide the required gas generator speed to drive the

accessories.

Ground Idle Speed is scheduled as a function of ambienttemperature.

− FLIGHT IDLE THRUSTIn flight operation, the FADEC will command the engine to Flight IdleThrust, which is scheduled in order to:− Avoid engine flameout, overtemperature or inability to accelerate.− Provide the required bleed airflow pressure and temperature for

the ECS and for the Anti-Icing System. If the FADECs receive anindication that the anti-icing system is on, Flight Idle thrust isrescheduled in order to provide the required air bleed flow,pressure and temperature. This automatic A-ICE Flight Idlerescheduling is inhibited below 15000 ft if the landing gear is downand locked.

− Enable the FADEC to accelerate the engine from Flight IdleThrust to 100% of the GO-around thrust mode in 8 seconds orless, at or below 9500 ft.

CLOSED-LOOP FAN SPEED CONTROL

The primary control mode of the engine is closed-loop fan speedcontrol. The fan speed requested by thrust lever is compared to themeasured fan speed. An error signal proportional to the differencebetween the request and measured fan speed is used to adjust thecommanded fuel flow to the engine to drive the fan speed error to zero.


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N1/N2 OVERSPEED/UNDERSPEED PROTECTION

The FADEC limits fuel flow to prevent the excessive rotor speed onboth the low-pressure rotor (N1) and the high-pressure rotor (N2). Ifthe fuel flow commanded by the closed-loop results in the surpassingof established rotor speed limits, fuel flow will be limited to that valuewhich will result in rotor speed limit.

The FADEC also incorporates a logic to initiate an engine shutdown ifthe upper limits of N1 and N2 are exceeded, in order to avoid apotentially destructive overspeed condition.

Logic within the FADEC incorporates a high-pressure rotor (N2)underspeed shutdown. This logic prevents damaging the turbine via anovertemperature condition if the engine attempts to operate at sub-idlespeed. If N2 drops below 54% the FADEC will command a shutdown.

The maximum steady-state rotor speeds are 100% N1 or 102.5% N2,(103.7% N2 for A1E engines). There is no minimum N1 speed.

INTERSTAGE-TURBINE TEMPERATURE (ITT) LIMITING

The FADEC has provisions for limiting engine fuel flow to preventexceeding ITT limits. If the fuel flow commanded by the closed-loopfan speed control exceeds established ITT limits, the FADEC will limitthe fuel flow to that value that will result in operation within the ITT limit.

ACCELERATION/DECELERATION LIMITING

Acceleration and deceleration limits within the FADEC logic restrict therate of commanded engine fuel flow to prevent surge duringacceleration or lean blow out during deceleration.

FLAMEOUT DETECTION/AUTORELIGHT

Flameout and autorelight detection logic within the FADEC detects anengine flameout and attempts an automatic relight before the engineloses power, if N2 is higher than 53%. In the event that a relight cannotbe successfully executed, the FADEC commands an engine shutdown.

During in-flight restarts, both ignition systems are energized.

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N1 REVERSIONARY CONTROL MODE

The FADEC provides a reversionary control mode to accommodate atotal loss of fan-speed (N1) signal.

The FADEC stores data on the correlation between N1 and N2 of anaverage engine in its non-volatile memory, and in the event that all N1signals are lost, it will control thrust governing N2 speed.

The engine control system is capable of modulating thrust in responseto thrust lever movement in the reversionary control mode. However,transient response times may be greater, minimum thrust may exceedflight idle thrust and maximum thrust may be less than that expectedduring normal control operation.

This mode is evident to the pilot due to the absence of N1 indication onthe EICAS.

FADEC INPUTS SELECTION AND FAULTACCOMMODATION

For every FADEC input, there is a selection and fault accommodationlogic, based on the inputs to both FADECs of the same engine.

The engine control system is highly fault tolerant. Because ofredundant sensor inputs and outputs, the control system canaccommodate multiple faults with no degradation in engine response.The fault accommodation philosophy is to maintain operation on thecontrolling FADEC for as long as possible before transferring control tothe standby FADEC.

For every detectable fault, the FADEC provides a signal to the EICASfor the alerting message and/or to the Central Maintenance Computerfor the maintenance message.

FADEC DISCRETE OUTPUTS

Each FADEC provides two discrete output signals, as follows:− N2 Speed Switch - Each FADEC activates a discrete output

whenever the engine is assumed to be running, based on N2.This signal is activated whenever N2 reaches (accelerating)56.4% and is deactivated whenever N2 drops below 53%.

− ECS OFF signal.


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ALTERNATE FADEC SELECTION

AUTOMATIC SELECTION− Whenever the FADEC in control is unable to safely control the

engine, it signals the alternate FADEC to automatically take overengine control.

MANUAL SELECTION− The alternate FADEC may be manually selected to control the

engine, by momentarily setting the FADEC Control Knob,located on the overhead panel, in the ALTN position.

The FADEC that is in control (A or B) is indicated on the EICAS.

FADEC RESET

The FADEC may be reset through the FADEC Control Knob. Uponreceiving the FADEC Control Knob input, the FADEC clears recordedinactive faults (faults not currently being detected).

In case any fault persists after the RESET command, it is not cleared.

Reset does not mean electrical power interruption to the FADEC.

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ENGINE OPERATIONGENERALThe Rolls-Royce AE 3007 engine uses an electronic control systembased on two Full Authority Digital Electronic Controls (FADECs) thatcontrol the engine. These FADECs interface with the engine, airframeand flight deck. A complete description of the engine control systemwas presented in the previous chapter.THRUST RATINGSThe engine control system schedules the corrected fan speed as afunction of pressure altitude, Mach number, ambient temperature, anti-ice system condition, thrust mode and thrust lever angle to achieve therated thrust conditions.Thrust ratings for the AE 3007 engine are:

Engines A1P A1EThrustratings Selectable ATTCS Selectable ATTCS

E TakeoffReserve

- - - E T/ORSV*

E Takeoff - - E T/O* E T/ORSV*

TakeoffReserve

- T/ORSV*

- T/ORSV*

Takeoff T/O* T/ORSV*

T/O* T/ORSV*

MaximumTakeoff-1 - - - -

AlternateTakeoff-1 ALT T/O-1* T/O-1* ALT T/O-1* T/O-1*

MaximumContinuous CON - CON -

E MaximumCllimb - - E CLB -

MaximumClimb CLB - CLB -

MaximumCruise CRZ - CRZ -

*Restricted to 5 minutesFor A1E engines, E T/O RSV and T/O RSV modes ate not intended fornormal operation. Their use must be recorded in the maintenancelogbook.For the respective takeoff rating, altitude, and Mach-number condition,fan speed is controlled to maintain constant thrust at any givenambient temperature below the flat-rated ambient temperature.


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ENGINE CONTROL

The engine control system controls the operation of the enginethroughout its operating envelope. The system modulates the fuel flowrate to the engine and the position of the variable geometry vanes(CVG) in response to inputs from the aircraft’s sensors andmeasurements of engine operating conditions. The engine controlsystem will not command a fuel flow that would result in exceedingrotor speed or temperature operating limits.

The engine control system is designed in such a manner that a singleelectrical failure will not cause significant thrust changes, result in anuncommanded engine shutdown or prevent a commanded engineshutdown. In case of loss of both FADECs, the engine control systemwill shut off fuel flow and move the CVGs to the closed position.

The engine control system performs two categories of functions: thrustmanagement and engine control. Thrust management logic interfaceswith the airframe and schedules a corrected thrust based on air dataand cockpit inputs. The fan speed request is passed to the enginecontrol logic, which controls the engine fuel flow and CompressorVariable Geometry (CVG) in response to the measured parameters inorder to attain the required engine response.

THRUST MANAGEMENT

This section of the FADEC software is responsible for functions directlyinvolved in the required thrust computation and management logic.Thrust management logic is provided to reduce flight crew workloadand enhance the aircraft’s operation.Thrust management functions are as follows: thrust mode selection,fan speed (N1) scheduling, Automatic Takeoff Thrust Control(ATTCS), Takeoff Data Setting (TDS), and thrust reverser interlock.

THRUST MODE SELECTION

Thrust logic management includes several thrust-rating modes that arecontrolled through associated buttons on the cockpit, set during thetakeoff data setting procedure, automatically triggered by the ATTCSor by advancing the Thrust Lever Angle (TLA) above the thrust setposition.Thrust-rating mode defines the available engine thrust at the existingambient conditions. The following thrust modes are available:

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ALTERNATE TAKEOFF (ALT T/O-1)

− All engines: This mode is the normal all engines operating takeoff mode andis available only through the use of the Takeoff Data Settingprocedure. Selection of this mode ensures the best engine durability andeconomy of operation. In this mode the ATTCS is active, so thatT/O-1 mode is triggered in case of engine failure.

MAXIMUM TAKEOFF-1 (T/O-1)

− A1P engine: This is the One Engine Inoperative (OEI) mode for the normal,all engines operating, ALT T/O-1 mode. In addition to beingselected by an ATTCS trigger, it may also be selected fromALT T/O-1 mode, at or below 1700 ft above takeoff pressurealtitude, by pushing the T/O thrust-rating button. It is not anormal pilot selectable takeoff mode.

− A1E engine: This is the One Engine Inoperative (OEI) mode for the normal,all engines operating, ALT T/O-1 mode. The FADECs will selectT/O-1 mode if the T/O switch is pressed and the current mode isALT T/O-1 during takeoff phase, if the ATTCS is triggered andthe current mode is ALT T/O-1 or if the thrust lever is movedbeyond Thrust Set position and the current mode is ALT T/O-1.


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TAKEOFF (T/O)

− A1P engine: This mode is the maximum, all engines operating takeoff mode.For engine durability and economy of operation, this modeshould only be selected when ALT T/O-1 is not authorized. ATTCS is active in this mode, so that ATTCS triggers upondetection of an engine failure, commanding a thrust increase toT/O RSV mode. The T/O mode is automatically selected atFADEC power up, and at the initialization of the Takeoff DataSetting procedure. T/O is also automatically selected in flightbelow or descending through 15000 ft provided the landing gearis down and locked. T/O is selected if there is weight on wheels,the TLA is at 50° or less and the T/O thrust-rating button ispushed. This mode is also selected when the T/O thrust-ratingbutton is pushed and the pressure altitude is greater than 1700 ftabove takeoff altitude.

− A1E engine: This is a medium thrust level, selectable through the TakeoffData Setting procedure, for all engines operating. For enginedurability and economy this mode should be selected ifconditions do not permit use of ALT T/O-1 but do not requireE T/O mode.

EXTENDED TAKEOFF (E T/O)

− A1E engine: This mode is the highest level, all engines operating, takeoffmode. For engine durability and economy of operation, thismode should only be selected when T/O mode is not authorized.In case of engine failure the ATTCS triggers the E T/O RSVmode. The E T/O is automatically selected at FADEC power-upand also at initiation of the Takeoff Data Setting procedure. ET/O is also automatically selected in flight, at or below 15000 ft,when the landing gear down and locked is received by theFADECs on both engines. This mode is also selected when theT/O button is pushed and the pressure altitude is greater than1700 ft above takeoff altitude. The FADECs will select E T/Omode if the T/O switch is pressed after takeoff phase, if the T/Oswitch is pressed and the current mode is T/O-1 or if the thrustlever is moved beyond Thrust Set position in flight or aftertakeoff phase.

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TAKEOFF RESERVE (T/O RSV)

− A1P engine:This mode is the corresponding OEI mode for all enginesoperating in T/O mode. The engine will produce the maximumrated thrust for the existing ambient conditions in this mode.T/O RSV is automatically selected when ATTCS is triggeredduring operation in T/O mode. T/O RSV is also selected if bothengines do not agree on the thrust mode or when the thrustmode of the remote engine cannot be determined. This modewill also be selected from the T/O mode, at or below 1700 ftabove takeoff altitude, when the T/O thrust-rating button ispushed.

NOTE: T/O RSV is manually selected by advancing one or bothTLA above Thrust Set position, regardless of any modepreviously selected.

− A1E engine:This is the corresponding OEI mode for all engines operating inT/O mode. This mode is accessible through a FADEC commandin response to an ATTCS triggering event. The FADECs willselect T/O RSV mode if the T/O switch is pressed and thecurrent mode is T/O during takeoff phase, if the ATTCS istriggered and the current mode is T/O or if the thrust lever ismoved beyond Thrust Set position and the current mode isT/O. This mode is also accessible by pressing the takeoff buttonwhile in T/O and the aircraft is in post takeoff condition or on theground.

NOTE: The use of this mode requires a notation in the aircraftmaintenance log.

EXTENDED TAKEOFF RESERVE (E T/O RSV)

− A1E engine:This mode is the corresponding OEI mode for all enginesoperating in E T/O mode. E T/O RSV is automatically selectedwhen ATTCS is triggered during operation in the E T/O mode.The FADECs will select E T/O RSV mode if the T/O switch ispressed and the current mode is E T/O or T/O RSV duringtakeoff phase, if the ATTCS is triggered and the current mode isE T/O, if the thrust lever is moved beyond Thrust Set positionand the current mode is E T/O or if the thrust lever is movedbeyond the Thrust Set position and the takeoff button is pressed. Use of this mode requires a notation in the aircraft maintenancelog.


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MAXIMUM CONTINUOUS (CON)

− All engines: This mode is selected by pushing the CON push button. CONmode is available when the pressure altitude is greater than300 ft above takeoff altitude and there is no landing gear downand locked, or when the pressure altitude is greater than 1700 ftabove takeoff altitude. The CON mode switch inputs to theFADECs are inhibited on ground.

MAXIMUM CLIMB (CLB)

− All engines: This mode is selected by pushing the CLB push button. CLBmode is enabled when the pressure altitude is greater than500 ft above takeoff altitude, there is no landing gear down andlocked signal and there is no OEI signal, or when pressurealtitude is greater than 1700 ft above takeoff altitude and there isno OEI signal. The CLB mode switch inputs to the FADECs areinhibited on ground. For A1E engines CLB is the default modewhen T/O or ALT T/O-1 is selected for takeoff.

EXTENDED CLIMB (E CLB)

− A1E engine: This mode is enabled under the same CLB conditions describedabove. However, E CLB is the default mode when E T/O isselected. Pressing the CLB button while in CLB mode togglesthe climb thrust to E CLB and vice-versa.

MAXIMUM CRUISE (CRZ)

− All engines: This mode is selected by pushing the CRZ push button. CRZmode is enabled when the pressure altitude is greater than500 ft above takeoff altitude, there is no landing gear down andlocked signal, and there is no OEI signal, or when pressurealtitude is greater than 1700 ft above takeoff altitude and there isno OEI signal.

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AE3007A1E THRUST MODE SELECTION

Thrust mode selection on A1E engines is a bit more complex than onthe other engines. The following tables illustrate how the thrust modescan be selected by pressing the T/O button, by advancing ThrustLevers above thrust Set or by the ATTCS.

PRESSING TAKEOFF BUTTON

Current Mode During takeoff phase (1) Post takeoff phaseALT T/O-1 T/O-1 E T/O

T/O-1 E T/O E T/OT/O T/O RSV E T/O

T/O RSV E T/O RSV E T/O (2)E T/O E T/O RSV E T/O

(1) Takeoff phase is configured when altitude is less than 1700 ftabove takeoff altitude, five minutes or less time has been elapsedsince thrust set selection for takeoff and current thrust mode is oneof the takeoff modes.

(2) T/O RSV to E T/O is a thrust decrease.

(3) If current thrust is E T/O RSV, flight altitude is between 1700 ftabove takeoff altitude and 15000 ft and the takeoff button ispressed, thrust will decrease to E T/O.

ADVANCING THRUST LEVERS ABOVE THRUST SET POSITION

Thrust Lever Angle above Thrust Set (TLA>78°)ATTCS NOT triggered

Current Mode During takeoff phase Post takeoff phase

ALT T/O-1 T/O-1 E T/O

T/O T/O RSV E T/O

E T/O E T/O RSV E T/O

CON, CLB, E CLBCRZ

- E T/O

T/O-1 (1) T/O-1 E T/O

T/O RSV (1) T/O RSV E T/O RSV

E T/O RSV (1) E T/O RSV E T/O RSV

(1) If the ATTCS is not triggered, these three modes are onlyaccessible by pressing the T/O button after selecting normal enginetakeoff modes through the Takeoff Data Setting procedure.


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Thrust Lever Angle above Thrust Set (TLA>78°)ATTCS triggered

TLA>78° andT/O button

Current Mode After ATTCS trigger TLA > 78° pressedALT T/O-1 T/O-1 T/O-1 E T/O RSV

T/O T/O RSV T/O RSV E T/O RSVE T/O E T/O RSV E T/O RSV E T/O RSV

Pushing the Takeoff Button with the Thrust Lever above Thrust Set willselect E T/O RSV mode regardless of the current takeoff mode or flightphase.

FAN-SPEED SCHEDULING

The thrust management logic calculates the corrected fan-speedrequest at any point in the flight envelope. The scheduled, correctedfan speed is computed as a function of pressure altitude, Machnumber, air temperature and other aircraft signals.

The thrust lever quadrant has five significant thrust positions definedas:

Thrust Lever Position Thrust Level Angle Maximum Reverse 0-4° Minimum reverse 14-22°

Idle 22-28° Thrust Set 72-78°

Maximum Thrust 78-85° Maximum reverse and maximum thrust are defined by mechanicalstops at either extremes of the thrust lever movement. Idle is definedby a mechanical gate that must be lifted to allow the trust lever totransition from forward flight to the reverse flight region. The thrust setposition on the thrust lever is delineated by a detent at 75°. For anygiven pressure altitude, Mach number and air temperature the FADECcomputes a corrected fan speed corresponding to the thrust leverposition. The fan speed computed for the thrust lever position isdependent upon the selectable thrust mode. The Target Thrust (N1Target) is defined as the thrust corresponding to the corrected fanspeed scheduled with the thrust lever at the Thrust Set position. Atarget thrust is computed for each thrust mode. Flight idle thrustcorresponds to the corrected fan speed with the TL at the idle positionand is independent of the thrust mode. The FADEC schedules thecorrected fan speed as a function of the thrust lever angle and thethrust mode to result in the linear relationships:

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A1P Engines

Any movement of the thrust levers above the Thrust Set positionresults in the scheduling of the maximum takeoff thrust, regardless ofthe current thrust mode except for A1E engines (refer to A1E ThrustMode Selection). A thrust lever position below the idle gate schedulesreverse thrust provided such action is enabled by the thrust reverserinterlock logic.


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ALTERNATE TAKEOFF THRUST CONTROL SYSTEM

During a takeoff, if an engine failure is detected, the ATTCSautomatically resets thrust on the remaining engine from AlternateTakeoff thrust to Maximum Takeoff thrust. In addition, depending ontakeoff thrust setting and ambient conditions, the FADECs generate anECS OFF signal to close the Pack Valves. (Refer to ECU operation onSection 2-14).

ATTCS ARMING CONDITIONS

ATTCS is armed when:

− Both engines are ATTCS capable,− Associated thrust lever angle is equal to or higher than 45°.

NOTE: ATTCS capable is defined as E T/O (A1E engine), T/O(A1P engine) or ALT T/O-1 (All engines) mode selected,with the airplane on ground and the engine running.

ATTCS TRIGGERING CONDITIONS

After being armed, the ATTCS is triggered under any of the followingconditions:

− The thrust lever for the opposite engine is reduced to below 38°TLA.

− Either FADEC for the on-side engine receives an oppositeengine or on-side engine inoperative condition, or a ThrustLever Angle limited to idle signal.

− The opposite engine does not indicate ATTCS being armed,within 2 seconds after the on-side engine ATTCS has armed.

− The opposite engine disarms ATTCS and the on-side enginedoes not disarm within 2 seconds.

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If ATTCS is armed and either FADEC A or B detects an oppositeengine inoperative condition, the controlling FADEC commands theon-side engine to a higher takeoff thrust, as shown in the table:

Engine Takeoff Selection

Two EnginesOperation

ATTCS TriggeredOne EngineOperation

A1P T/O T/O RSV

ALT T/O-1 T/O-1

ALT T/O-1 T/O-1

A1E T/O T/O RSV

E T/O E T/O RSV

ATTCS DISARMING CONDITIONS

The ATTCS disarms if any of the following conditions is met:− After being armed, the Thrust Lever Angle is reduced below 42°.− ATTCS is triggered on either engine.− No ATTCS capable takeoff mode is selected.

NOTE: If thrust lever is moved beyond the THRUST SET position theFADEC automatically commands the engine to the maximumavailable thrust, disregarding the takeoff mode selected, exceptfor A1E engine (see A1E engine Thrust Mode Selectionsection).

TAKEOFF DATA SETTING

The Takeoff Data Setting function is provided in order to enable thepilot to input reference data into the FADEC prior to takeoff. Such datawill be used to calculate N1TARGET during takeoff. The following datahas to be inputted:

− Takeoff Mode (T/O MODE), which corresponds to:− T/O or ALT T/O-1 (A1P engine).− E T/O, T/O or ALT T/O-1 (A1E engine).

− Reference Takeoff Temperature (REF TO TEMP), which shallcorrespond to the Static Air Temperature (SAT) on the groundprovided by the Air Traffic Control Tower, ATIS (AutomaticTerminal Information Service) or other accurate source.


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− Reference Takeoff Anti-Ice Condition (REF A-ICE), which is theanti-ice system condition (ON/OFF) that the FADEC willconsider to calculate N1TARGET.

This function is enabled during ground operations only and with thrustlever angle below 50°, before or after engine start.

The takeoff data setting is performed through the Takeoff Data Settingcontrols (STORE button and SET control) on the overhead panel.

After selecting the takeoff page on the MFD, The Takeoff Data Settingprocedure shall be as follows:

a) After the first pressing of the STORE button, the MFD indicates thefollowing initial values for the three takeoff data:

− T/O MODE: T/O for A1P engine; E T/O for A1E engine.

− REF TO TEMP: T2SYN (if engine is running) or ISA Temperature (otherwise).

NOTE: - T2SYN is the synthesized total air temperature at theengine fan inlet.

- T2.5 is the fan discharge total air temperature.

− REF A-ICE: OFF.

An arrow points to T/O MODE line. Through the SET Control thetakeoff mode ALT T/O-1 may be selected.

b) At the second pressing of the STORE button, the arrow points toREF TO TEMP, indicating that this parameter may be adjusted.Through the SET control, the initial value may be adjusted to therequired temperature. Each momentary command of the SETcontrol will increase (INC) or decrease (DEC) the current value by1°C. If the SET control is held at the command position for morethan 1 second, the REF TO TEMP is changed by 5°C/sec.

NOTE: The acceptable REF TO TEMP value range is limited toT2SYN ± 10°C.

c) At the third pressing of the STORE button, the arrow points toREF A-ICE line, indicating that this parameter may be adjusted.Through the SET control, the initial condition (OFF) can be switchedto ON and back to OFF alternately.

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d) At the fourth pressing of the STORE button:− If the engines are running and the REF TO TEMP is within limits

(T2SYN ± 10°C):− The FADECs accept the takeoff data and successfully

terminate the procedure.− The MFD displays the takeoff data.− The FADEC begins to calculate and display the N1TARGET

based on the takeoff data.− If the engines are not running, the adjusted takeoff data will

remain displayed in amber color, which means that they havenot been accepted yet. Then:− After engines start, if the adjusted REF TO TEMP is within

limits, the FADECs accept the takeoff data and successfullyterminate the procedure, the MFD displays the takeoff data,and the FADEC begins to calculate and display theN1TARGET based on the takeoff data.

− Otherwise, the takeoff data will not be accepted by theFADECs and the MFD will display dashed lines for all takeoffdata in amber color, and a caution message (ENG NO TODATA) is presented on the EICAS if TLA > 45°.

− In order to enter the correct takeoff data, the procedure mustbe started again, through the STORE button.

e) If, after takeoff data had been successfully entered, the pilot needsto correct any of them, the STORE button must be commandedagain in order to restart the procedure.

f) In case of disagreement between the REF A-ICE condition selected

by the pilot and the actual Anti-Ice system condition, a cautionmessage (ENG REF A/I DISAG) is displayed on the EICAS,provided the Parking Brake is released (OFF) or with any ThrustLever Angle above 45°.

g) If any thrust lever is set to an angle above 45° before takeoff datasuccessfully entered, a caution message (ENG NO TO DATA) ispresented on the EICAS.


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ENGINE START

Engine start, commanded through the Start/Stop Knob, is automaticallymanaged by the FADEC as follows:

− The FADECs A and B alternate as FADEC in control on everysubsequent ground start, if the Ignition Selector Knob is set toAUTO position, as a single ignition system, corresponding to theFADEC in control, will be used.

− The FADEC activates the ignition system when N2 is atapproximately 14% and commands the fuel solenoid valve toopen when N2 is at approximately 31.5% (28.5% for airplanesequipped with FADEC B7.4 and on) or 12 seconds after ignitionis activated, if the Ignition Selector Knob is set to AUTO or ONposition.

− Whenever the start cycle is completed, the FADEC deactivatesthe ignition system and provides a discrete signal to commandthe Starting Control Valve (SCV) to close.

− If the Ignition Selector Knob is set to OFF position, the FADECneither activates the ignition system nor actuates the fuel valvefrom closed to open position, in order to enable ground/flight drymotoring.

NOTE: If the engine is already running with TLA above IDLEthrust, the fuel valve is not closed, even if the IgnitionSelector Knob is set to OFF position.

− The FADEC monitors Interturbine Temperature (ITT) start limitoverride during ground starts. If the temperature exceeds thecontrol temperature reference, the FADEC reduces fuel flow.There is no automatic engine shutdown by the FADEC for anovertemperature on start. When the engine is started on ground,only the FADEC in control commands ignition, if the IgnitionSelector Knob is set to AUTO position. During an in flight start,both FADECs command ignition.

− If a flameout is detected, the FADEC turns on the ignitionsystem, provided the ignition switch is in the AUTO position,until the engine is restarted.

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ENGINE DRY MOTORING

An Engine Dry Motoring must be performed for at least 30 secondsafter any aborted start to assure that no unburned fuel remains in thecombustion chamber and/or to reduce residual ITT prior to attemptinganother start.Ignition switch must be rotated to Off position in order to disableignition and fuel flow prior to rotating the Stop/Run/Start switch ti thestart position.

ENGINE SHUTDOWN

Normal engine shutdown, through the Start/Stop Knob, is managed bythe FADEC, which commands the engine fuel solenoid valve to close.The normal sequence only occurs with the thrust levers positioned atIdle. Thrust levers should be positioned at IDLE before the Start/Stopknob is positioned at Stop.A shutdown sequence is also performed whenever N2 is below 54%.

NOTE: The Engine Fire Extinguishing Handle, when actuated, alsoshuts the engine down by closing the respective fuel shutoffvalve, interrupting fuel supply from the wing tanks.


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ATTCS FAIL ATTCS failure associated witha low N1.

E1 (2) ATTCS NO MRGN The engine has no ITT or N2margin to achieve higherthrust if ATTCS is trigged.

WARNING

E1 (2) OIL LOW PRESS Oil pressure has droppedbelow 34 psi and the engine isrunning or the pressure switchhas failed at the closedposition and the engine is notrunning.

ENG 1 - 2 OUT N2 from both the engines hasdropped below 8500 rpm(underspeed shutdown limit)uncommanded

E1 (2) LOW N1 Engine does not achieverequestd N1.

E1 (2) FUEL LO TEMP The fuel temperature in theengine has dropped below 5°C.

E1 (2) ATS SOV OPN The engine ATS shutoff valve(SCV) remained open above53% N2.

CAUTION

ENG REF A/I DISAG Disagreement between theREF A-ICE condition selectedby the pilot and the actual anti-icing system condition hasbeen detected by the enginecontrol associated withParking Brake released (OFF)or with any TLA above 45°.

E1 (2) CTL FAIL A failure in the Engine controlsystem has been detected.

ENG1 (2) TLA FAIL Thrust Lever Angle sensor hasfailed.

ENG NO TO DATA Takeoff Data has not beensuccessfully entered withengine running and above53% N2.

(Continued)

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TYPE MESSAGE MEANINGFADEC ID NO DISP There are different FADEC

applications installed in theaircraft.

ENG 1 (2) OUT N2 has dropped below 8500rpm (underspeed shutdownlimit) uncommanded.

CAUTION E1 (2) EXCEEDANCE ITT or N2 exceeded thecurrent ITT or N2 limit duringan interval of the flight leg.

E1 (2) FPMU NO DISP An incompatible FPMU wasinstalled on a A1E engine.

E1(2) NO DISP Associated FADEC hasdetected a non-dispatch failurecondition.

E1 (2) OIL IMP BYP The differential pressureacross the oil filter hasexceeded the normal range.

E1 (2) FADEC FAULT A dispatchable MMELcategory B FADEC fault wasdetected.

ADVISORY

E1 (2) FUEL IMP BYP The differential pressureacross the fuel filter hasexceeded the normal range.

E1 (2) SHORT DISP A dispatchable MMELcategory B FADEC fault wasdetected.

CHECK XXX PERF(XXX=A1P,A1E)

Inform the FADEC applicationinstalled in the aircraft.Displayed only on ground withflaps 0° and parking brakesapplied.


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CONTROL PEDESTAL

1 - GUST LOCK LEVER

Limits thrust lever movement and locks the elevator controlsurfaces when set in LOCKED positionRefer to Section 2-13 − Flight Controls.

2 - THRUST LEVER

MAX - Provides maximum takeoff thrust.THRUST SET - Provides N1TARGET thrust setting.IDLE - Provides ground and flight idle thrust settings.MAX REV - Provides maximum reverse thrust.

NOTE: Protection against inadvertent thrust reverser command inflight is provided through the mechanical idle stop and theelectrical flight idle stop.

3 - FRICTION LOCK

Rotated clockwise, thrust lever movement becomes progressivelymore resistant, so that thrust levers will not slip.

4 - THRUST RATING MODE buttons

T/O - Selects maximum takeoff thrust-rating mode.CON - Selects maximum continuous thrust-rating mode.CLB - Selects maximum climb thrust-rating mode.CRZ - Selects maximum cruise thrust-rating mode.


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CONTROL PEDESTAL

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POWERPLANT CONTROL PANEL1 - IGNITION SELECTOR KNOB

OFF - Deenergizes the ignition system.AUTO - FADECs control the ignition system automatically,

depending on the engine requirement.ON - Commands the FADEC to activate continuously the two

ignition channels.

2 - FADEC CONTROL KNOB (SPRING-LOADED TO NEUTRAL)RESET - Resets the FADECs, and clears faults.ALTN - Alternates the FADEC in control.

NOTE: The knob becomes inoperative if held in any position formore than 3 seconds.

3 - TAKEOFF DATA STORE BUTTON− Initiates and terminates takeoff data setting.− At the first pressing, an arrow points to T/O MODE line.− At the second pressing allows REF TO TEMP adjustment.− At the third pressing allows REF A-ICE to be input.− At the fourth pressing, if REF TO TEMP is within limits, the

takeoff data is accepted and the procedure is successfullyaccomplished.

− For complete procedures refer to Takeoff Data Settingparagraph.

NOTE: The button becomes inoperative if held pressed for morethan 3 seconds.

4 - TAKEOFF DATA SET CONTROL− When turned, selects the T/O MODE, increases (INC) or

decreases (DEC) the REF TO TEMP value and also switchesthe A-ICE condition state presented on the MFD during takeoffdata setting.

− Momentary actuation changes the REF TO TEMP values by1°C. If the control is held for more than 1 second at the INC orDEC position, REF TO TEMP is charged by 5°C/sec.

− The mode T/O can be switched to ALT T/O-1 and back to T/Oalternately (A1P engine).

− The mode E T/O, T/O and ALT T/O-1 can be switchedalternately (A1E engine).

− The A-ICE initial condition (OFF) can be switched to ON andback to OFF alternately.


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5 - START/STOP SELECTOR KNOB

STOP - Commands the FADEC to shut the engine down, providedassociated Thrust Lever is at IDLE.

RUN - Allows normal engine operation.START - This is a momentary position that initiates the engine start

cycle. If the knob is held in this position for more than 3seconds, it becomes inoperative. In this case, a FADECreset command is required.

NOTE: The Start/Stop selector knobs are equipped with atransparent protection guard over the knob for better engineidentification.

POWERPLANT CONTROL PANEL

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FIRE HANDLE

The Fire Handle, located on the Fire Protection Control Panel, allowsengine emergency shutdown. For further information on fire controls,refer to Section 2-07 − Fire Protection.

ENGINE INDICATION ON EICAS

1 - N1TARGET INDICATION

− Corresponds to the maximum available engine thrust for a giventhrust-rating mode, airspeed, ambient condition, and bleed airstatus.

− Digits are cyan.− Ranges from 0 to 100% RPM with a resolution of 0.1%.− Indicated by a cyan T-shaped bug.− Indication is removed from the display for request values greater

than 100% or less than 0%.

2 - THRUST-RATING MODE ANNUNCIATION

− Indicates the current thrust-rating mode.− Labels: T/O or ALT T/O-1 (A1P engine);

E T/O, T/O or ALT T/O-1 (A1E engine); CON, CLB, or CRZ.− Color: cyan.− When engines operate in alternate takeoff mode a green

ATTCS annunciation is presented below the takeoff label toindicate that the ATTCS system is armed.

3 - THRUST REVERSER ANNUNCIATION

− Indicates the position of the upper and lower Thrust Reverserdoors.

− Label: REV.− Color:

− Fully open: green.− In transition: amber (if applicable).


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4 - N1 INDICATION

− Displays N1 in RPM percentage.− Scale:

− Ranges from 0 to 100%. Extends up to 110% if exceedingthe red line.

− Colors: green from 0 to 99.9%. red line at 99.9%.

− Digits:− Ranges from 0 to 120% RPM, with a resolution of 0.1%.− Colors: green from 0 to 99.9%.

red at 100.0% and above.

5 - FADEC IN CONTROL ANNUNCIATION

− Indicates the FADEC channel that is controlling the engine.− Labels: A or B.− Color: green.

6 - IGNITION CHANNEL ANNUNCIATION

− Indicates the ignition channel that is enabled.− Labels: IGN A, IGN B, IGN AB, or IGN OFF.− Color: green.

7 - INTERTURBINE TEMPERATURE INDICATION

− Scale:

− During engine start:− green from 300 to 800°C.− red line at 801°C.

− Takeoff mode:− green from 300 to 947°C (A1P engine).

from 300 to 992°C (A1E engine).

− red line at 948°C (A1P engine). at 993°C (A1E engine).

− CON, CLB and CRZ modes:− green: from 300 to 900°C (A1P engine).

from 300 to 935°C (A1E engine). − amber: from 901 to 947°C (A1P engine). from 936 to 970°C (A1E engine). − red line at 948°C (A1P engine). at 971°C (A1E engine).

− If the red line is exceeded, the scale extends a further 50°C.

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− Digits:− Ranges from -65 to 1999°C with a resolution of 1°C.− Color: corresponds to the color of the scale.

8 - N2 INDICATION

− Displays N2 in RPM percentage.− Digits:

− Ranges from 0 to 120% RPM with a resolution of 0.1%.− Colors:

A1P engines:− green from 0 to 102.4%.− red line at 102.5%.

A1E engines and FADEC B7.6 and before:− green from 0 to 103.8%.− red line at 103.9%.

A1E engines and FADEC B8:− green from 0 to 105%.− red line at 105.1%.

9 - FUEL FLOW INDICATION

− Ranges from 0 to 2000 KPH (or 4000 PPH) with a resolution of5 KPH (or 10 PPH).

− Color: green.

10 - LOW-PRESSURE AND HIGH-PRESSURE TURBINE VIBRATION INDICATION

− Ranges from 0 to 2.5 inches per second (IPS).− Low-pressure scale and pointer colors:

− green from 0 to 1.8 IPS.− amber above 1.8 IPS.

− High-pressure scale and pointer colors:− green from 0 to 1.1 IPS.− amber above 1.1 IPS.

11 - OIL TEMPERATURE INDICATION

− Ranges from 0 to 180°C with a resolution of 1°C.− Scale, pointer, and digit colors:

− amber below 21°C.− green from 21 to 126°C.− red above 126°C.


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12 - N1 REQUEST BUG

− Indicates N1 requested by the Thrust Lever position.− Indicated by a green filled triangle.− Ranges from 0 to 100% RPM.− Indication is removed from the display for request values greater

than 100% or less than 0%.

13 - OIL PRESSURE INDICATION

Scale, pointer, and digit colors depend on the FADEC version asshown below:

(1) For N2 < 88% the amber band between 34 psi and 50 psi doesnot exist, and the green band lower limit is 34 psi.

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ENGINE INDICATION ON EICAS


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TAKEOFF PAGE ON MFD 1 - TAKEOFF MODE INDICATION

− Indicates Takeoff Mode as selected through the Takeoff DataSet Control.

− Label: T/O or ALT T/O-1 (A1P engine). E T/O, T/O or ALT T/O-1 (A1E engine).− In flight, the indication is removed from the display.

2 - REFERENCE TAKEOFF TEMPERATURE INDICATION− Indicates reference takeoff temperature as adjusted through the

takeoff data set control.− In flight, the indication is removed from the display.

3 - REFERENCE ANTI-ICE STATUS INDICATION− Indicates reference anti-ice status as selected through the

takeoff data set control.− Labels: ON or OFF.− In flight, the indication is removed from the display.

4 - OIL LEVEL INDICATION− Ranges from 0 to 13 US Quarts for left engine and from 0 to 14

US Quarts for right engine with a resolution of 1 US Quart.− Digits:

− green from 6 to 14 US Quarts.− amber below 6 US Quarts.NOTE: The right engine is capable of measuring a higher oil

level due to sensor position.

TAKEOFF PAGE ON MFD

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FIRST ENGINE BACKUP PAGE ON RMU

− Contains thrust modes, N1, ITT, N2, Fuel Flow, Oil Pressure andOil Temperature indications.

− Only the N1 indication contains analog and digital indication. Theother indications are in digital format.

− Label and legend color: white.− Data color limits: same as the EICAS display.

FIRST ENGINE BACKUP PAGE ON RMU


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THRUST REVERSER

GENERAL

Each engine is equipped with thrust reverser.

The thrust reverser is for ground operation only, and its function is todirect engine exhaust gases forward and outwards to producedeceleration of the airplane.

The thrust reverser system consists of an electric control/indication, anhydro-mechanical actuation system, and two pivoting doors.

When stowed, the thrust reverser is part of the exhaust nozzle.

LOCK PROTECTION

The system incorporates three locking systems to avoid inadvertent in-flight deployment. The actuators and doors are mechanically locked inthe stowed position through the primary and secondary locks. In casethe primary and secondary reverser locks fail, the tertiary lock preventsthe door from deploying. In the stowed position, the doors are held bythe primary lock only, with the secondary and tertiary locks remainingunloaded. The primary and secondary locks are electricallycommanded/controlled and hydraulically powered to unlock. Thetertiary lock is electrically commanded/controlled and electricallypowered to unlock, thus providing a separate and fully independentlocking system.

OPERATION

The thrust reverser is commanded by the backward movement of theThrust Lever. Upon selection, the mechanical locks are removed andhydraulic pressure is applied to deploy the thrust reverser doors. Inreverser mode, the doors rotate about a fixed axis. Rotation of thedoors is controlled by extension and retraction of the hydraulic dooractuators.

After pivoting, the rearmost part of the doors blocks the normal nacelledischarge path and directs the flow through the aperture created by itsrotation.

The loss of electrical and/or hydraulic power does not result ininadvertent deployment.


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OPERATION LOGIC

Each FADEC will command Maximum Reverse thrust on ground only,when the associated thrust reverser is deployed and associated thrustlever is requesting reverse thrust whenever either of the followingconditions are met:- Airplane on the ground indication from both main landing gears, and

main landing gear wheels running above 25 kt, or- Airplane on the ground indication from both main landing gears and

from nose landing gear.

During landing, when the Thrust Levers are set to below IDLE, theFADEC commands reverse thrust only after the Thrust Reverser doors(both engines) are completely deployed. If the Thrust Lever isrequesting forward thrust, the FADEC will command IDLE thrust if theassociated engine thrust reverser indicates that there is a “not stowed”or a “deployed” condition.

If one engine is inoperative or one thrust reverser is not deployed, theFADEC of the operative side will only command Reverse Thrust if theassociated Thrust Lever is requesting reverse thrust and the ThrustLever of the affected side is set to IDLE. Such a feature is provided toavoid uncommanded thrust asymmetry.

EICAS INDICATION

An indication of right and left thrust reversers deployed is presented onthe EICAS. If a failure or a disagreement is detected, a cautionmessage is presented on the EICAS.

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THRUST REVERSER INTERLOCK

The FADECs interface with the thrust reverser system of thecorresponding engine.

Each FADEC receives two pieces of information from the thrustreverser system:− Stowed: If all doors of the corresponding engine are stowed.− Deployed: If all doors of the corresponding engine are deployed.

For flight operation there is also a flat between IDLE andMAX REVERSE position. The FADEC enables reverse thrustdepending on the position of the reverser doors and on the position ofthe engine thrust lever, and reduces the engine thrust to IDLE, if thereis an indication of an inadvertent thrust reverser deployment in flight,which normally is not possible due to the Flight Idle electrical stop.

EICAS MESSAGES


ENG1 (2) REV FAIL

-Thrust reverser doorsnot stowed and in transitwith Thrust Levers set ator above IDLE, or

-Thrust Levers set belowIDLE in flight.

CAUTION

ENG1 (2) REV DISAGREE

-At least one thrustreverser door not fullyopen, or

-Thrust reverser systemnot isolated fromhydraulic system (ThrustLever set at or aboveIDLE), or

-Door locking or positionswitch signal failure withThrust Levers set at orabove IDLE (groundonly).

ADVISORY E1 (2) IDL STP FAIL Idle stop has failed.


MANUAL

2-10-70 Page

4 Code



AIRPLANEOPERATIONS

MANUAL

POWERPLANT


5 Code

01

THRUST REVERSER


MANUAL

2-10-70 Page

6 Code



Embraer Emb-135bj Aom Vol. 2 Aom

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