This document is to be used for training purposes only. Under no circumstances it supercedes or replaces any official information published by the manufacturer.
The information contained herein is proprietary to GE Commercial Finance Aviation Training (GECAT) and is disclosed in confidence. It is the property of GECAT, and shall not be reproduced or disclosed in whole or in part, or used for any purpose whatsoever without the express written consent of GECAT.
Issue: Jan07
Issue: Jan07 Issue: Feb02 Revision: 00
FOR TRAINING ONLY GE Capital Aviation Training. Reproduction Prohibited
Chapter 27-00
Page 2
190 Abbreviations and parts locations
190 A ABC ABS AC ACC ACE ACMP ACOC ACARS ADA ADC ADF ADG ADS ADSP AEO AETC AFCS AFM AGB AGCU AGT AH AICC A/I A/I AIP AIOP ALC ALF ALT AMS Issue: June06 Revision: 00
Av. Bus Control Absolute Value Alternating Current Aft Core Cowl/ Active Clearance Control Actuator Control Electronics AC Motor Pump control Air Cooled Oil Cooler Airborne Communication Adressing & Recording System Air Data Application Module Air Data Computer Automatic Direction Finder Air Driven Generator Air Data System Air Data Smart Probe All Engines Operating AC Essential Transfer Contactor Automatic Flight Control System Airplane Flight Manual Accessory Gearbox Auxiliary Generator Control Unit Agent Amp Hour Auxiliary Integrated Control Centre Anti-Ice Approach Idle Autonomous Input Processor Actuated Input Output Processor Auxiliary Generator Line Contactor Aft Looking Forward Altitude Air Management System
AMJ AMLCD AMM AMS AN A/O AOA APM APR APU ARINC ARP AS ASC ASCB ASV A/T ATA ATC ATE ATS ATTCS ATTD ATOA AUX AUX GEN AWC
Advisory Material Joint Active Matrix Liquid Crystal Display Aircraft Maintenance Manual Air Management System Aerospace Force-Navy Air/ Oil Angle of Attack Air Pressure Module Automatic Power Reserve Auxiliary power unit Aeronautical Radio Incorporated Aerospace Recommended Practice Aerospace Standard APU Start Contactor Avionics Standard Communication Bus Anti-Surge Valve Auto Throttle Air Transport Association Air Traffic Control Automatic Test Equipment Air Turbine Starter Automatic Takeoff Thrust Control System Attendant Above Take-off Altitude Auxiliary Auxiliary Generator Aural Warning Computer
B B BARO
FOR TRAINING ONLY Reproduction Prohibited
Ball Barometric Setting Chapter 0-00
Page 1
BARO-ALT BATT BC BCD BCM BCS BCV BFE BIC BIC BIT BITE BLWR BNR BPT BRG BRK BTC BTL BTMS
Barometric Altitude Battery Battery Contactor Binary Coded Decimal Brake Control Module Brake Control System Brake Control Valve Buyer Furnished Equipment Backplane Interface Controller Bus Interface Controller Built-In Test Built-in Test Equipment Blower Binary Numeric Representation Break Power Transfer Bearing Brake Bus Tie Contactor Bottle Brake Temperature Monitoring System
C C CAN CAS CAWS CB CBM CBP CCA CCD CCD Issue: June06 Revision: 00
Centigrade/ Celsius Control Area Network Crew Alerting System Central Aural Warning System Circuit Breaker Circuit Breaker Module Circuit Breaker Panel Control Card Assembly Cursor Control Device Compliance Check Database
CCDL cc/h CCP CCPS CCS CCT CCW CDP CDU CF CFC CFE CFR CFSP CG CH CLB CLB-1 CLB-2 CMC CMF CMM COM CON CONFIG CPCI CPCS CPU CRC CRES CRG CRI CRT FOR TRAINING ONLY Reproduction Prohibited
Chross Channel Datalink cubic centimeter per hour Cockpit Control Panel Cockpit Control Position Semsors Cabin Communications System Cockpit Control Transducer Counter clockwise Compressor Discharge Pressure Control Display Unit Commercial Fan Carbon Fiber Composite Customer Furnished Equipment Code of Federal Regulations (USA) Cargo Fire Suppression Panel Centre of Gravity Channel Climb Climb 1 rating Climb 2 rating Central Maintenance Computer Communications Management Function Component Maintenance Manual Communications Maximum Continuous rating Configuration Computer Program Configuration Item Cabin Pressure Control System Central Processing Unit Cyclic Redundancy Check Corrosion Resisting Steel Cargo Certification Review Item Cathode Ray Tube Chapter 0-00
Page 2
190 CRZ CSD CT CTA CTA CTR CVFC CVR CW
Cruise rating Constant Speed Drive Current Transformer Current Transformer Assembly Centro Tecnico Aerospacial Centre Cargo Vent Fan Contactor Cockpit Voice Recorder Clockwise
DMU DMU DOC DP DP3(190) DPDT DRH DRL DU DVDR DWLK
Drier/Metering Unit Data loader Management Unit Document Differential Current Bleed Bias Sensor Dole Pole Double Throw Dual responder Heritage Dual Responders (revision L) Display Unit Digital Voice and Data Recorder Downlock
D Da DADC DAU dB DB DC DCU DCPC DCTC DDG DEOS DET DFDAU D/I DISAG DISC D/LNA DMC DME DMM Issue: June06 Revision: 00
Double amplitude Digital Air Data Computer Data Acquisition Unit decibel Database Direct Current, electrical Direction Control Unit DC Power Centre DC Tie Contactor Dispatch Deviation Guide Digital Engine Operating System Detector Digital Flight Data Acquisition Unit Descent Idle Disagree Disconnected Diplexer/Low Noise Amplifier Direct Maintenance Cost Distance Measuring Equipment Data Memory Module
E E1 E2 EBAY EBU EC ECS ECU ECP EDP EDS EED EEPROM EGPWS EGT EHCL EHSV EICAS EICC FOR TRAINING ONLY Reproduction Prohibited
Engine 1 Engine 2 Electronic Bay Engine Built Unit / Engine Built-up Essential Contactor Environmental Control System Electronic Control Unit Engine Configuration Plug Engine Driven Pump Electronic Display Systems Electro-explosive Device Electronically Erasable Programmable Read Only Memory Enhanced Ground Proximity Warning System Exhaust Gas Temperature Electro Hydraulic Cowl Lock Electro Hydraulic Servo Valve Engine Indicating and Crew Alerting System Essential (Emergency) Integrated Control Centre Chapter 0-00
Essential Integrated Control Document Embraer Liebherr Equipamentos do Brasil Empresa Brasileira de Aeronautica Electromagnetic Compability Emergency Electro-Magnetic Interference Engine External Power AC Contactor Emergency Parking Brake System Emergency Parking Brake Valve External Power DC Contactor Electrical Power Generating and Distribution System Erasable Programmable Logic Device External Power Mode Erasable Programmable Read Only Memory Embraer Regional Jet Electronic Starter Control Essential (Power bus) Essential Transfer Contactor Engine Thrust Compensation Electronic Thrust Trim system Engine Vibration Monitor Engine Vibration Control System External Power
F F FAA FADEC FAN FAP FAR Issue: June06 Revision: 00
Fahrenheit Federal Aviation Administration Full Authority Digital Engine Control (Electronic Controller) Fan Flight Attendant Panel Federal Aviation Regulations (U.S.A.)
Fly By Wire Flight Control Module Flight Control System (primary and secondary) Fuel Conditioning Unit Flush Control Unit Flow Control Valve Fuel Cooled Oil Cooler Flight Data Recorder Forward Flight Attendant Panel Flight Hour Functional Hazard Assessment Flight Idle First In First Out Fire Extinguisher Flight Flexible take-off Flight Management Computer Flight Management System Fuel Metering Unit Fuel Metering Valve Fuel/Oil Foreign Object Damage Fuel Overhead Panel Forward Outer Seal Field Programmable Gate Array Fuel Quantity Gauging System Fuel Quantity Processor Fuselage Station Full Scale Deflection Full Scale Output Feet Fault Tree Analysis File Transfer Protocol Chapter 0-00
Page 4
190 FWD FWSOV
Forward Fire-wall Shut-off Valve
G GA GA RSV GCR GCS GCU GE GE/ GEAE GEN G/I GLC GMAP GMO GND GP GPM GPS GPU GS GS GSE GSTC
Go-around rating Go-around reserve Generator Control Relay Generator Control Switch Generator Control Unit General Electric Company General Electric Aircraft Engines Generator Ground Idle Generator Line Contactor Ground Mapping Ground Maintenance Override Ground Guidance/Display Control Panel US Gallons per Minute Global Positioning System Ground Power Unit Glideslope Ground Spoilers Ground Support Equipment Ground Service Transfer Contactor
I
High Definition Data Link Control High Efficiency Particulate High Frequency
IAC IAS ICC ICD ICU ID
H HDLC HEPA HF Issue: June06 Revision: 00
HIRF HOR HP HP HP HPA HPC HPC (190) HPRSOV HPSOV HPT HPTCC (190) HPX Hr HRD HSI HSP HTR HTS HW HX/ Hex HYD Hz
FOR TRAINING ONLY Reproduction Prohibited
High Intensity Radiated Fields Hold Open Rod Horsepower High Pressure High Pass High Power Amplifier Hydraulic Pump Contactor High Pressure Compressor High Pressure Regulation Valve High Pressure Shut-Off Valve High Pressure Turbine High Pressure Turbine Clearance Control Valve Horsepower Extraction Hour High Rate Discharge Horizontal Situation Indicator Hydraulic Synoptic Page Heater High Thermal Stability Hardware Heat Exchanger Hydraulic Hertz (Cycles per second)
Integrated Avionics Computer Indicated Air Speed Integrated Control Circuit Interface Control Document Isolation Control Unit Identification Key Chapter 0-00
Page 5
IDG IEVM IFS IFSD IGA IGB (190) IGV ILS IM IMP INBD Inco inHG I/O IOM I/P IPSA IPT IPV (190) IRS ISA ISO ITT
Integrated Drive Generator Integrated Engine Vibration Monitor Inner Fixed Structure (Forward Core Cowl) Inflight Shutdown Intermediate Gain Antenna Inlet Gearbox Inlet Guide Vane Instrument Landing System Inner Marker Impending Inboard Inconel Inches of Mercury Input/Output Input/Output Module Input Integrated Pitot Static Angle of Attack Probe Inadvertent Parallel Trip Inlet Pressurizing Valve Inertial Reference System International Standard Atmosphere International Standards Organization Inter-Turbine Temperature
J JAA JAR JTSO
Joint Aviation Authorities Joint Aviation Regulations (Europe) JAA Technical Standard Order
K KG Issue: June06 Revision: 00
Kilogram
kHz km kohm KPa KPad KPag kph KVA kW
Kilo Hertz Kilometers kilo-ohm Kilo-Pascal Kilo-Pascal delta Kilo-Pascal gage kilograms per hour Kilo Volt Amp Kilowatt
L L LAN LAV LB/ lbs/hr Lbs/Min Lbs/sec lbf LCD LDG LG ECU LGS LHE L/I LICC LLP LLS LLV LMS LOC LOP FOR TRAINING ONLY Reproduction Prohibited
Liter Local Area Network Lavatory Lbs Pound pounds per hour pounds per minute pounds per second Pounds-force Liquid Crystal Display Landing (Gear) Landing Gear Electronic Control Unit Landing Gear System Left Hand Engine Lamding Idle Left Integrated Control Centre Life Limited Part Liquid Level Sensor Low Limit Valve Load Management System Localizer Low Oil Pressure Chapter 0-00
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190 LOTC LP LPT LRD (190) LRM LRU LSB LSS LVDT
Loss of Thrust Control Low Pressure Low Pressure Turbine Low Rate Discharge Line Replaceable Module Line Replaceable Unit Low Significant Bit Lightning Sensor System Linear Variable Differential Transducer
M m mA MAU Max mB MCDU MCOEI MCU MCU MEL MES MF MFD MFP MFS Mhz MIC Mils MIN MKR BCN Issue: June06 Revision: 00
Meter milli-amperes Modular Avionics Unit Maximum milli Bar Multi-Function Control Display Unit Max Continuous One Engine Inoperative Modular Concept Unit Motor Controller Unit Minimum Equipment List Main Engine Start Multi-Function Multi-Function Display Multi-Function Probe Multi-Function Spoiler Megahertz Microphone Measuring unit. equivalent to 1/1000 inch Minutes Marker Beacon
MLG MLI MLW MM MMF mm Mn MON MPD MPR MPU MRC MS ms msec MSR mV
Main Landing Gear Magnetic Level Indicators Maximum Landing Weight Middle Marker Monitor Warning Function Millimeter Mach Number Monitor Material and Processes Directives Manual Power Reserve Magnetic Pickup Unit Modular Radio Cabinet Military Specification millisecond millisecond Motor Starting Relay Millivolt
N N Rotation Speed N/A Not Applicable NACA National Advisory Committee for Aeronautics NAI Nacelle Anti Ice NAPRSOV Nacelle Pressure Regulating and Shut off Valve NAS National Aerospace Standard NAV Navigation NB Narrow Band NBPT No Break Power Transfer NC No Change NCD No Computed Data NDOT Core SpeedAcceleration Rate FOR TRAINING ONLY Reproduction Prohibited
Chapter 0-00
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NIC NIM NLG No. NTO NVM N/W NWS NWSCM N1 N2 N2S
Network Interface Controller/ Computer Network Interface Module Nose Landing Gear Number Normal Takeoff Non Volatile Memory Nose Wheel Nose Wheel Steering Nose Wheel Steering Control Module Engine Low Pressure Rotor Speed/ Physical Fan Speed Engine High Pressure Rotor Speed/ Physical Core Speed Corrected High Pressure Rotor Speed
O OAT OBV OC OD ODS OEI OF OFV OGV OM O/P OTC OUTBD OV OVHT
Outside Air Temperature Operational (Operability) Bleed Valve Overcurrent Outside Diameter Overheat Detection System One Engine Inoperative Overfrequency Outflow Valve Outlet Guide Vane Outer Marker Output Outer Torque Coupling Outboard Overvoltage Overheat
P P Issue: June06 Revision: 00
Pressure Port
P0 P-ACE PACIC PAL PAST PAX PBA PBA P-BIT PBVA PCA Pcd PCU PDCSM pC/g PDD PDU 1PDT PFD PFTO PGE PI-BIT PLA PLD PM PMA PMG PMP PMS P/N PO POR POS FOR TRAINING ONLY Reproduction Prohibited
Ambient Pressure Power Actuator Control Electronic Passenger Adress & Cabin Interphone System Programmable Array Logic Pilot Activated Self Test Passenger Push Button Annunciated Push Button Actuator Periodic Built In Test Parking Brake Valve Applied Power COntrol Actuator Pressure compressor discharge Power Control Unit Parameter Dispenser& Collector State machine pico Coulombs per unit acceleration Periodic Device Driver Power Drive Unit One Pole Double Throw Primary Flight Display Power For Take Off Page Key Pilot Initiated Built In Test Power Lever Angle (TLA) Programmable Logic Device Processing Module Permanent Magnet Alternator Permanent Magnet Generator Pump Permanent Magnet Starter Part number Ambient Pressure Point of Regulation Position Chapter 0-00
Power-on Self Test pounds per hour Power ready Processor Protection Pressure Regulating and Shutt off Valve Phase Sequence Static Pressure Compressor Discharge Pressure Proximity Sensor Evaluation Module Power Supply Module Pounds per square inch, absolute Pounds per Spuare inch (psi) Pounds per square inch, gauge Proximity Sensing System Total Pressure Polytetrafluoroethylene Power Takeoff Assembly Power Transfer Unit Per Unit Power-up Built In Test Pulse-Width Modulated Power
R R RAM RAT RCB RBHA RDI Issue: June06 Revision: 00
Return Port/ Roller Random Access Memory Ram Air Turbine Radio Control Bus Regulamentos Brasileiros de Homologação Aeronáutica Refuel Defuel Indicator
Refuel Defuel Panel REFReference Radial Drive Shaft Reverser / Reverse Thrust 1 per Revolution Radio Frequency RAT Generator Control Unit Right Hand Engine Reverse Idle Right Integrated Control Centre RAT Line Contactor Load Resistance Radio Management Unit Output Resistance Read Only Memory Revolution Per Minute Radio System Bus Root Sum Square Source Resistance Reserve Radio Technical Commission for Aeronautics Resistive Thermal Device Ready to Load Rejected Take Off Rotary Variable Differential Transformer Reduced Vertical Separation Rearward
S S1, S2 S-ACE SAC
FOR TRAINING ONLY Reproduction Prohibited
Solenoid1, 2 Secondary Actuator Control Electronics Single Annular Combuster Chapter 0-00
Society of Automotive Engineers Satellite Communications Starter Air Valve Software Critical Level Signal Control Unit Starter Control Valve Set Delay System Description Source Destination Identifier System Definition Drawing Smoke Detection and Fire Suppression System Description Note Satellite Data Unit Second Service Synchronized Feedback Actuator Specific Fuel Consumption Shaft Horsepower Synchronized Locking Actuators Sea Level Static Status Matrix Smoke Shut Off Valve Solenoid Operated Valve Secondary Power Distribution Assembly Single Pole Double Throw Speaker Squelch Static Random Access Memory System Requirements Document Structural Repair Manual System Requirement Specifications System Safety Assessment
Static Source Error Correction Sign Status Matrix Solid State Power Controller Stand-by Stand-by Contactor Standard Day Store Key System Test Specification Service Servo Valve Switch Switch Software Switch, Pole A Switch, Pole B Stall Warning Protection SystemSYNCSynchronous Synchronous SystemTACThrust Asymmetry Compensation
T T2 TAC TACAN TAMB TAT TBC TBD TBV (190) TC TCAS TCPS TCQ TCS FOR TRAINING ONLY Reproduction Prohibited
Inlet Temperature Thrust Asymmetry Compensation Tactical Air Navigation Static Air Temperature Total Air Temperature To Be Calculated/Confirmed To be defined Transient Bleed Valve Thermocouple Traffic Alert and Collision Avoidance System Temperature Compensated Pressure Switch Thrust Control Quadrant Touch Control Stearing Chapter 0-00
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190 TDC TDS TGB(190) Ti TLA TLD TMS TO TO-1 TO-2 TO-3 TO/GA T/R T2 T3 T22 T25 (190) T45 TQA T/R TRAS TRF (190) TRS TRU TRUEC TRX TS TSO TST TVP
Issue: June06 Revision: 00
Top Dead Centre Takeoff Data Set Transfer Gearbox Titanium Throttle Lever Angle / Thrust Lever Angle Time Limited Dispatch Thrust Lever Angle Take-off Takeoff Mode 1 Takeoff Mode 2 Takeoff Mode 3 Take Off and Go Around Thrust Reverser Fan Total Inlet Temperature Compressor Discharge Temperature Compressor Inlet Temperature Compressor Inlet Temperature Inter-Turbine Temperature Throttle Quadrant Assembly Thrust Reverser Thrust Reverser Actuation System Turbine Rear Frame Thrust Rating Selector Transformer Rectifier Unit Transformer Rectifier Unit Essential Contactor Thrust Reverser Position Technical Specification Technical Standard Order Test Key True Vapor Pressure
U UF US G UTC UV
Underfrequency US Gallon Universal Time Coordination Undervoltage
V V V1 VAC VBPCI VBV (190) VDC VDR VDT VG VGV VGX VHF Vibe VIDL VNAV VOL VOR VREF VRMS VSCF VSCV VSV
FOR TRAINING ONLY Reproduction Prohibited
Volts the minimum speed in the takeoff Volts Alternating Current Virtual Backplane Peripheral Component Variable Bleed Valve Volts Direct Current VHF Data Radio Varialble Differential Transducer Variable Geometry Variable Guide Vane Variable Geometry Actuator Position Feedback Very High Frequency Vibration VOR/ILS Datalink Vertical Navigation Volume Very High Frequency Omnidirectional Radio Voltage Reference Volts Root Mean Square Variable Speed Constant Frequency Variable Speed Constant Frequency Variable Stator Vane
Wing Anti-Ice Fuel Flow Ratio Units: ratio of fuel flow to PS3 HMU Metering Valve Position Feedback Walter Kidde Aerospace Weight Off Wheels Weight On Wheels Weight On Wheels PSEM1 Weight On Wheels PSEM2 Warning Wheel Spin Wheel Spin Discrete Water Waste System Controller Weather
° °C °F > < +/-
Degrees Degrees Celsius Degrees Fahrenheit is greater than is less than plus or minus Ohm Differential
%
Percent
X XDCR XFER XFEED XPDR
Transducer Transfer Cross feed Transponder
Y YD
Yaw Damper
Z ZID
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Zone Isolation Device
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Chapter 0-00
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190
Intentionally Left Blank
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FOR TRAINING ONLY Reproduction Prohibited
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Figure 1: Cockpit Instruments Location
GP
PFD
Audio panel
MFD
MCDU
EICAS
CCD
Provision for future MCDU
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Chapter 0-00
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190 Figure 2 : Fwd E-Bay Equipment L WS WIPER
R WSWIPER
ICE-DETECTOR
ICE-DETECTOR
SPDA 1
MRC 1
PA ALTITUDE SENSOR
LSS
TAT TAT SMARTPROBE 1 SMARTPROBE 2 IRU 1
IRU 2 P-ACE 2 P-ACE 1
TCAS
DVDR BATT 1 Forward E-Bay
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Chapter 0-00
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Figure 3: Fwd E-Bay back
RH
LH
SMART-PROBE
GCU (RAT)
DIMMER
PRESSURE CONTROLLER
INVERTER MAU 2
MAU 1
EICC
Forward E-Bay
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Chapter 0-00
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190 Figure 4 : Center E-Bay
SF-ACE
MRC 2
SPDA 2
SF-ACE
MAU 3
FCU
MS
FIREX BOTTLES
LICC RICC
Center E-BAY
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Figure 5 : Aft E-Bay
ELT LIGHT
DVDR
HF ANT COUPLER
ELT- NAV P-ACE 3
HF SATCOM VHF VOR 3 DVDR
APU-FADEC
HS ACE WATER CTL AICC BATT 2
AFT E-BAY
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190
Intentionally left blank
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190 ATA 00 Documentation
190 Table of Content
EXTINGUISHING AND RESCUE MANUAL (IGFER). . . . . . . . . . 19 MAINTENANCE FACILITY AND
190 0-00 Documentation Introduction All maintenance checks, inspections, repairs, replacements and troubleshooting must be performed in accordance with valid documentation. The related documentation necessary to maintain the aircraft includes:
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Chapter 0-00
Page 1
Figure 1: Documentation
TECHNICAL PUBLICATION AIPC AMM ARM APM CMM CPC CPM FIM ITEM IGFER MFEP
Issue: June06 Revision: 00
Aircraft Illustrated Parts Catalog Aircraft Maintenance Manual Aircraft Recovery Manual Aircraft Planning Manual Component Maintenance Manual Consumable Products Catalog Corrosion Prevention Manual Fault Isolation Manual Illustrated Tool and Equipment Manual Instruction for Ground Fire Extinguishing and Rescue Maintenance Facility&Equipment Planning
TECHNICAL PUBLICATION Nondestructive Testing Manual Parts Information Letter Power Plant Buildup Manual Ramp Maintenance Manual Service / Information Bulletin Service News Letter (SNL) Standard Wiring Practices Manual Standards Manual Structural Repair Manual System Schematic Manual Task Card System Wiring Manual
REVISION Annually As required Quarterly Quarterly As required As required Semiannually Semiannually Quarterly Semiannually -
CUSTOMIZATION Always Not Applicable Not Applicable Always Not Applicable Not Applicable On request Not Applicable Always Always Always Always
Quarterly
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190 AMM PART I - SYSTEM DESCRIPTION SYSTEM (SDS) • Purpose: - Detailed description and explanation of the location, configuration, function, operation and control of the complete system (chapter), and its subsystems. - Enable the operator / mechanic / trainee to understand the three levels of overall construction, operation and function to the extent necessary to perform adequate maintenance and fault isolation of the system. • Available links: FIM, AIPC, ITEM, MPP, SM, SSM, WM. • Arrangement:
INTRODUCTION GENERAL DESCRIPTION
SDS Sections
COMPONENTS OPERATION TRAINING INFORMATION POINTS
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Figure 2: AMM Part 1 - SDS Example
AIRCRAFT MAINTENANCE MANUAL
AMM Part 1 - System Description Section (SDS) Example: IDG 2
RICC
RAT
APU GEN
RAT GCU EICC STATIC INVERTER
LICC
IDG 1
EM170SDS240043.DGN
Components GENERATOR DRIVE SYSTEM (24-21) The generator drive system is the usual source of aircraft electrical AC (Alternating Current) power in flight and on the ground during taxi and takeoff procedures. APU AC GENERATION (24-22) The APU (Auxiliary Power Unit) generation system is used primarily when the aircraft is on the ground for aircraft maintenance or flight preparation. This system can also be used to let the aircraft be dispatched with an altitude restriction or as a backup source of electric power in flight. EMERGENCY AC POWER GENERATION (24-23) The EPGDS (Electrical-Power Generation-and-Distribution System) is capable of generating electrical power in an emergency condition. This ensures that the essential loads will be supplied with power for an unlimited time if the main AC (Alternating Current) generators are lost. STATIC INVERTER (24-24) The Static Inverter converts aircraft 28 VDC (Volt Direct Current ) to single phase 115 VAC. The figure AC GENERATION - BLOCK DIAGRAM provides further data on the preceding text.
AC GENERATION - BLOCK DIAGRAM
24-20-00
EFFECTIVITY: ALL
Page 7 Dec 08/03
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190 AMM PART II - MAINTENANCE PRACTICES AND PROCEDURES (MPP) • Purpose: - Contains all necessary maintenance practice and procedure data to enable the mechanic to maintain the aircraft properly, at the level of line, hangar / service centre maintenance actions, or line ramp level. • Available links: - WM, AIPC, CMM, SWPM, CPM, FIM, MPP, SDS, SM, SRM, SSM, ITEM.
• Arrangement:
PAGEBLOCKS
COMPONENT MAINTENANCE LOCATION
PRACTICES
(100)
(200)
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SERVICING
(300)
REMOVAL
ADJUSTMENT
INSPECTION
CLEANING
INSTALLATION
TEST
CHECK
PAINTING
(400)
(500)
(600)
(700)
REPAIRS
(800)
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Figure 3: AMM Part II - MPP Example
AIRCRAFT MAINTENANCE MANUAL
INTEGRATED DRIVE GENERATOR (IDG) - REMOVAL/INSTALLATION 1. General A. This section gives the procedures to remove/install the IDGs. B. The IDG 1 is installed on the left engine gearbox and the IDG 2 is installed on the right engine gearbox. C. The procedures in this section are given in the sequence below. DESCRIPTION
190 MAINTENANCE PRACTICES AND PROCEDURES (MPP) The AMM provides sufficient information to enable a mechanic to service troubleshoot, test, adjust and repair systems and to remove and install any component on the line or in the hangar. This AMM is written in accordance with the ATA 100 Revision 28 recommendations. Revision 28 has, among other things, these two recommendations • Simplified English • Prepared for AMTOSS (Aircraft Maintenance Task Oriented Support System)
Aircraft Maintenance Task Oriented Support System (AMTOSS) AMTOSS eases maintenance procedures by giving the specific procedure (task) and the primary steps of procedures (sub tasks) a different number identification code. The identification code numbers have a minimum of five elements and a maximum of seven. The elements are: • An extension to the three ATA 100 Chapter, section and subject numbering. • After the first three elements a function code will follow. This code identifies the type of work that has to be done. • The fifth element shows the task or sub task identification number. This number is applicable to the task/sub task only and is counting the sequence of the task or sub task. Numbers 801 thru 999 are for tasks and 001 thru 800 for sub tasks. • The sixth element identifies the differences in configuration and the related procedures and techniques. • Issue: June06 Revision: 00
• Tasks/sub tasks that are special to the operator, or which are written by sub contractors are identified in the seventh element.
Function Codes The function codes used in AMTOSS are from 000 thru 900. Each code represents the following function: • • • • • • • • • •
Each of the codes represents the task. The codes are subdivided to identify the tasks or sub tasks following. The page below shows the application of the function codes 400. Further information is found in the AMM Introduction section.
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Figure 4: AMTOSS Codes and Subdivision
FUNCTION CODE (INSTALLATION)
ASSIGNED BY MANUFACTURER
ASSIGNED BY ATA (REF. 1-3-1)
SEQUENTIAL ORDER (INSTALLATION NO. 1)
EQUIPMENT IDENTIFIER
CONFIGURATION INSTALLATION No. 1 ON N0001-N9999
(ENG. FIRE EXTING. BOTTLES)
PRIMARY TASK NUMBERS
26 - 22 - 03 - 400 - 801 - B
ELEMENT NUMBERS
1
SUBTASK NUMBER (SEE NOTE)
26 - 22 - 03 - 42X - XXX - A01 - CC1
2
35
4
6
7
SYSTEM/CHAPTER (FIRE PROTECTION)
ASSIGNED BY OPERATOR
SUBSYSTEM/SECTION ( EXTINGUISHING.) NECESSARY TO IDENTIFY:
SUB-SUBSYSTEM (ENGINE FIRE EXTING.)
UNIT/SUBJECT (BOTTLES)
DIFFERENCES IN CONFIGURATIONS (SB) SHOWN AS AN ALPHA CHARACTER IN THE FIRST DIGIT DIFFERENCES IN METHODS/TECHNIQUES OF TASK ACCOMPLISHMENT SEQUENCE VARIATIONS OF STANDARD PRACTICES AT HARDWARE LEVEL MULTIPLE SHEETS OF AN ILLUSTRATION OR TABLE
FUNCTION CODE
CREATION OF SEQUENTIAL ORDER WITH THE SAME FUNCTION CODE WITHIN THE SUBJECT (TASK-801 TO 999 SUBTASK-001 TO 800)
RESERVED FOR MANUFACTURER'S DISCRETIONARY USE, FOR FURTHER DEFINITION OF THE FUNCTION
NOTE: SUBTASK ARE THE MAINTENANCE ACTIONS REQUIRED TO ACCOMPLISH A TASK. UNIQUE IDENTIFIERS FOR SUBTASK FUNCTIONS SHALL COMMENCE AT 001 AND TERMINATE AT 800.
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EM170MPP000007.DGN
(INSTALL UNIT)
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190 AIRCRAFT MAINTENANCE MANUAL An Effectivity box may have more information, e.g. “Collins”, to point out an item that differs in configuration.
Division of Subject Matter: The first section of the AMM is the “Introduction”. This section provides an introduction to the manual and instructions how to use the manual correctly. It is strongly recommended to read this section of the manual. The introduction section also includes a list of the chapters that are included in the manual. Each chapter has the following items filed at the front: • Effectivity Code Cross Reference List • Highlights page(s) for each revision • List of Effective Pages • List of Effective Temporary Revisions • Service Bulletins • Table of Content Effectivity: Effectivity Codes are used in the AMM to allow many operators with different configuration aircraft to correctly apply the manual to their aircraft. The effectivity is shown in the lower left hand corner of each page in the AMM. The following are the effectivities used: • MASTER • ALL OPERATORS • An operators prefix e.g. “LX” (Swiss) • An “E” Code (E + a number, e.g. E26-018
MASTER: Always means a “master” manual. It is possible for an operator to have a master manual. ALL: “All” is always shown together with an operator prefix or “ALL OPERATORS” in the box. Example: EFFECTIVITY: ALL LX In this case, it means that the page applies to all LX aircraft. EFFECTIVITY: ALL OPERATORS This page applies to all aircraft. E Codes: The E code is shown together with MASTER or an operator prefix. Example: EFFECTIVITY: E26-018 MASTER If a chapter has one or more pages with an E code, there will be a green “Effectivity code cross reference list” in the beginning of the chapter. This list identifies the aircraft serial numbers that are affected by the E code.
If a chapter has one or more pages with an E code, there will be a green “Effectivity code cross reference list” in the beginning of the chapter. This list identifies the aircraft serial numbers that are affected by the E-code.
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Figure 5: Aircraft Maintenance Manual
OPERATORS
LX
E 26- - 018
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190 FAULT ISOLATION MANUAL (FIM) • Purpose: - Provides all information needed to report and correct aircraft faults to avoid or reduce dispatch delays and fix defective items or systems. • Available links : AMM, WM, AIPC. • Arrangement:
INTRODUCTION
FIM
FIM USAGE
CAS MESSAGES LIST OBSERVED FAULTS LIST
FAULT REPORT SECTION
CABIN FAULTS LIST
FAULT ISOLATION SECTION
MAINTENANCE MESSAGES LIST CAS MESSAGES LIST OBSERVED FAULTS
CHAPTERS
CABIN FAULTS FAULT CODE INDEX MAINTENANCE MESSAGES INDEX FAULT ISOLATION PROCEDURES TASK SUPPORT DATA
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Figure 6: Fault Isolation Manual used with aircraft CMC (Central Maintenance Computer)
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190 RAMP MAINTENANCE MANUAL (RMM)
COMPONENT MAINTENANCE MANUAL (CMM)
• Purpose:
• Purpose :
- Provides information which can improve the ground handling and avoid delays when difficulties are encountered for the dispatchability, at ramp level.
- Provides information and procedures applicable to a workshop environment for the return of a component to a serviceable condition. • Arrangement:
• Arrangement :
DESCRIPTION AND OPERATION TESTING AND FAULT ISOLATION
SPECIAL PROCEDURES REMOVAL INSTALLATION SERVICING STORAGE REWORK APPENDICES
Chapter 0-00
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Figure 7: Ramp Maintenance Manual Example
RAMP MAINTENANCE MANUAL
ZONE 252
B
A ZONES 123 124
BATTERY 1
BATTERY 2
C
C
A B
C TYPICAL
Main Batteries (24-36-00) Figure 2
EFFECTIVITY: ALL
COMPONENT LOCATION
24-30-00 Part 1
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190 WIRING MANUAL (WM)
SYSTEM SCHEMATIC MANUAL (SSM)
• Purpose:
• Purpose:
- Provides the necessary information concerning the wiring diagrams to enable fault isolation and maintenance.
- Provides technical information to aid the maintenance personnel in understanding the aircraft systems and performing the fault isolation procedures at the LRU (Line Replaceable Unit) level. - The information is presented through diagrams, with indication of component location, system interface, and references to other manuals (SSM, SDS, MPP and WM).
• Arrangement :
INTRODUCTION
• Available links : SSM-Internal link, SDS MPP, WM.
CHAPTERS 21 TO 80 – WIRING DIAGRAMS
• Arrangement:
WM
CHAPTER 91 – ELECTRICAL AND ELECTRONIC CHARTS AND LISTS INTRODUCTION
SSM
FIRST LEVEL - SYSTEM BLOCK DIAGRAMS SECTIONS
SECOND LEVEL - SYSTEM SCHEMATIC DIAGRAMS THIRD LEVEL - SYSTEM LOGIC DIAGRAMS
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Figure 8: Wiring Manual and System Schematic Manual AIRCRAFT WIRING MANUAL
190 AIRCRAFT ILLUSTRATED PARTS CATALOGUE (AIPC) General The AIPC is provided by the manufacturer for use in provisioning, requisitioning, storing and issuing replaceable parts and units, and for identifying parts. The AIPC is a companion document to the AMM and includes all parts for which maintenance practice has been provided. Section Numbering The section numbering is made of three elements, whereby the first and second element represents the chapter/section breakdown according ATA 100. To enable quick location of installation figures and to simplify the task of locating items within the IPC, the third element in the Chapter numbering is designated to aircraft major zones as follos: • • • • • •
00 Electrical installations in all applicable zones 01 Forward fuselage and cockpit 02 Center fuselage and cabin 03 Aft fuselage and cargo compartment 04 Wings 05 Engine and nacelles
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Figure 9: Aircraft Illustrated Parts Catalogue (AIPC) AIRCRAFT ILLUSTRATED PARTS CATALOG
INTEGRATED DRIVE GENERATORS 1/2 (IDG) INSTL, AC GENERATION 1/2, LH/RH NACELLE þ ITEM NOT ILLUSTRATED
24-21-01
Fig. 01 Page 1 Mar 29/04
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190 OTHER DOCUMENTATION
CORROSION PREVENTION MANUAL (CPM) • Purpose:
AIRCRAFT RECOVERY MANUAL (ARM)
- Provides information on materials and procedures for prevention and removal of corrosion damage to aircraft as well as to display EMBRAER’s recommendations about frequent corrosion troubles.
• Purpose: - Contains information in sufficient detail to effect recovery in the most expeditious manner while maintaining consideration of recovery personnel safety and prevention of additional damage to the aircraft.
ILLUSTRATED TOOL AND EQUIPMENT MANUAL (ITEM) • Purpose:
AIRPORT PLANNING MANUAL (APM)
- Provides all information about GSE (Ground Support Equipment) to support the operation and maintenance of the aircraft and all its onboard equipment.
• Purpose: - Provides necessary information to enable a proper planning of the airports for the aircraft operation.
INSTRUCTIONS FOR GROUND FIRE EXTINGUISHING AND RESCUE MANUAL (IGFER) CONSUMABLE PRODUCTS CATALOGUE (CPC)
• Purpose: - To provide the necessary information to guide ground rescue teams while rescuing passengers in case of aircraft accidents.
• Purpose: - Provides the information about the consumable materials used to overhaul and repair the aircraft.
Lifting the Aircraft with Jacks Figure 3-30-05-1 Sheet 1
EFFECTIVITY: ALL
3-30-05 Page 4 Feb 10/04
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190 MAINTENANCE FACILITY AND EQUIPMENT PLANNING (MFEP)
SERVICE BULLETIN (SB) • Purpose: - Presents modifications or special inspections to be carried out on in-service aircraft.
• Purpose: - Provides aircraft characteristics to assist airline personnel responsible for service, line maintenance, overhaul facilities and equipment planning. - It identifies and describes aircraft maintenance and operational facility requirements.
• Arrangement: • - Planning information: involves those aspects related to plan the aircraft maintenance opportunity/necessity for the SB incorporation. - Material: information to assist the operator in obtaining the material necessary for the SB incorporation. - Accomplishment instructions: presents the step-by-step instruc tions and illustrations for accomplishing the work.
NONDESTRUCTIVE TESTING MANUAL (NDT) • Purpose:
INFORMATION BULLETIN (IB)
- Provides all general procedures of nondestructive tests acceptable by Embraer for investigating the quality and integrity of materials and components.
SERVICE BULLETIN (SB) • Purpose: - Presents modifications or special inspections to be carried out on in-service aircraft. • Arrangement: Issue: June06 Revision: 00
• Purpose: - It is used to transmit information, which are not related to actions requiring a record of accomplishment. • Arrangement: - Purpose - Effectivity (applicability) - Description - Approval - References - Affected Publications
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Figure 11: Example of Nondestructive Testing Manual NONDESTRUCTIVE TESTING MANUAL
Internal Detailed Visual-Inspection of the Direct Vision Stops, Tracks, Roller Studs and Bellcrancks Figure 1
56-10-00 Part 5
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w-ndt1803
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190 STRUCTURAL REPAIR MANUAL (SRM) • Purpose: - To permit the operators to identify and evaluate the damage and restore the structural integrity of the aircraft by means of a repair or by replacing the damage part.
Available links: MPP, NDT, CPM, SM, ITEM, CPC.
STANDARD WIRING PRACTICES MANUAL (SWPM) • Purpose: - Allows the operator to repair, manufacture and handle all harness-related components.
TASK CARD SYSTEM (TCS) • Purpose: - Provides a reliable tasks list that can be customized by operators to support the MPD and suit the operators needs.
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Figure 12: Example of Structural Repair Manual STRUCTURAL REPAIR MANUAL
A
EM170SRM520058.DGN
DENT
A Dent on Forward Passenger Door Skin - Location Figure 101
190 04-00 Aircraft General Introduction The following short introduction will give you a general overview of the Embraer 190-100 and 190-200.The Emb 190 is a low wing, twin engine jet airplane of conventional structure, designed for medium to short range operations. The Emb 190-100 has a total length of 36.15 meters (118 ft 7 in.), a wing span of 28.56 meters (93 ft 8 in.) and an approximate height of 10.48 meters (34 ft 5 in.). The EMB 190-200 has a total length of 38-67 (126 ft 10 in), a wing span of 28.72 m (94 ft 3 in) and an approximate height of 10.55 m (34 ft 7 in). It also features a pressurized cabin. The fuselage has a so-called double bubble design.
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Figure 1: The Embraer 190-100(Emb 190)
Embraer 190-100 Emb 190
34 ft 5 in (10.48 m)
118 ft 7 in (36.15 m)
93 ft 8 in (28.56 m) 32 ft 10 in (10.00 m)
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190 The fuselage and conditioning packs The fuselage is pressurized between the forward pressure bulkhead, located forward of the cockpit, and the aft pressure bulkhead, which is located behind the aft electronics bay. Normal pressurization control is automatic, and the conditioned air is provided by two air conditioning packs located in the unpressurized area forward of the wing root.
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Figure 2: The fuselage and conditioning packs
Fwd pressure bulk head
Aft pressure bulk head
The conditioned air is provided by two air conditioning packs located in the unpressurized area forward of the wing root.
Normal pressurization control is automatic
Conditioning packs
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190 Cockpit and cabin sections The cockpit can accommodate 2 crew members in the pilot seats and one observer on the jump seat. Two flight attendants. The Embraer 190-100 is designed for 98 passengers and the Embraer 190-200 is designed for 108 passengers. There are 2 Galleys and 2 toilets - one of each in the front and the aft sections of the cabin. The cabin also features a wardrobe, built next to the front passenger entrance. There are 2 cargo compartments below the passenger cabin - one in front and one behind the wing fairing.
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Figure 3: Cockpit and cabin sections
Two galleys and two toilets
One jump seat
Two pilots seats Wardrobe 98 passenger seats
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190 The landing gear The Emb 190 has a forward retracting twin-wheel nose landing gear (NLG). The NLG has a normal steering angle of about 76°, making the aircraft highly manoeuvrable. The shock absorbers are of oleo-pneumatic type. The steering motor, taxi light and one landing light are mounted on the NLG. When the NLG is extended, the rear doors remain open while the front doors reclose after extension or retraction. The main landing gears (MLG) also have oleo-pneumatic shock absorbers and twin wheels retract sideways.
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Figure 4: The landing gear
EMBRAER 190
The NLG has a normal steering angle of about 76°, making the aircraft highly manuverable.
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The main landing gears (MLG) retract sideways.
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190 Emb 190 turning radius With a full deflection of the nose wheel, the Emb190-100 can theoretically turn on taxiways as narrow as 21.40 m (70 ft 3 in.). The Emb 190-200 can turn on taxiways as narrow as 22.68 m (74 ft 5 in). Note that the largest clearance is required by the tail, which is not visible. In a maximum turn, either the left or the right MLG remains stationary, marking the centre of the turn.
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Figure 5: Emb 190-100 turning radius
Nosewheel 14.07 m (46 ft 2 in.) Nosetip 18.12 m (59 ft 5 in.)
Wingtip 18.39 m (60 ft 4 in.)
Tail 20.51 m (67 ft 4 in.)
With a full deflection of the nose wheel, the Emb190-100 can theoretically turn on taxiways as narrow as 21.40 m (70 ft 3 in.).
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Notes:
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Figure 6: Emb 190-200 turning radius
Nosewheel 15.10 m (49 ft 7 in.) Nosetip 19.13 m (62 ft 9 in.)
Wingtip 18.61 m (61 ft 1 in.)
Tail 21.9 m (71 ft 10 in.)
With a full deflection of the nose wheel, the Emb190-200 can theoretically turn on taxiways as narrow as 22.68 m (74 ft 5 in.).
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190 CF34-10E high bypass turbofan engines The two wing-mounted CF34-10E high bypass turbofan engines are based on the CF34 engine family, which is widely used in aviation. Engine controls and fuel scheduling are provided by a full-authority digital engine control (FADEC) with fully modular design. The CF34-10E incorporates the aerodynamic efficiency of wide cord fan, which produces most of the engine's 18,500 Lbs maximum thrust. To enhance aircraft braking capability, the fan by-pass air is reversed not the engine.
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Figure 6: CF34-10E high bypass turbofan engines
CF34-10E high bypass turbofan engine
To enhance aircraft braking capability, the fan by-pass air is reversed. Side View
Top View
FADEC
Vent to Compartment RHS Cooling Scoop
FADEC
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190 The Embraer Emb 190-100 The Emb 190-100 can climb to 30,000 ft within 15 minutes, and has a certified ceiling of 41,000 ft. The Emb 190-100 has a maximum cruising speed of mach.80 Depending on the long or normal range version, it can reach destinations upwards of 2,000 nautical miles with standard reserves left in the tanks.
The Embraer Emb 190-200 The Emb 190-200 can climb to 30,000 ft within 17 minutes and has a certified ceiling of 41,000 ft. The Emb 190-200 has a maximum cruising speed of mach.80 Depending on the long or normal range version, it can reach destinations upwards of 2,000 nautical miles with standard reserves left in the tanks.
.
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Figure 7: The Embraer Emb 190
The Emb 190: Can climb to 30,000ft within 15 minutes Certified ceiling of 41,000 ft Maximum cruising speed of mach .80 Range up to 2,000 nautical miles
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190 Emb 190 composite structure To maximize performance, a variety of modern composites like fiberglass and carbon have been used. These materials are lighter in weight and more durable than conventional aluminium, improving aircraft performance. Besides the conventional flight controls, the aircraft is equipped with an adjustable stabilizer and multi functional spoilers. Aerodynamic characteristics are enhanced by leading edge slats and ground spoilers.
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Figure 8: ERJ 190 composite structure
Multi function spoilers
Ground spoilers
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190 04-01 Differences EMB 190-100 /190-200 The Embraer 190-100 has a seat capacity of 98 seats. The Embraer 190-200 has a seat capacity of 108 seats.
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Figure 1: Seat capacity
Embra e r 190-100 98 Seats
EMBRAER 190-100 98 seats MTOW : 47790 kg / 105357 lbs MLW : 43000 kg / 94797 lbs MZFW : 40800 kg / 89947 lbs
Embra e r 190-200 108 Seats
EMBRAER 190-200 108 seats MTOW : 48790 kg / 107562 lbs MLW : 45000 kg / 99206 lbs MZFW : 42500 kg / 93695 lbs
190 5-00 Time Limits and Maintenance checks Introduction Chapter 05, time limits and maintenance checks of the aircraft maintenance manual, provides inspection procedures for various scheduled and unscheduled checks. The data concerning detailed time limits and scheduled maintenance checks can be found in the "Maintenance Planning Guide".
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Figure 1: Chapter 05
Chapter 5:
States inspection procedures for various scheduled and unscheduled checks Issue: June06 Revision: 00
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190 Unscheduled maintenance checks Unscheduled maintenance checks have to be performed after the following occurrences: • • • • • • • • • • • • • • •
Lightning strike Hard landing or overweight landing High drag/side-load landing conditions Strong turbulence or buffeting conditions High-load-factor flight Landing-gear-down overspeed Exceeding flap/down speed condition Bird strike Engine fire warning or overheat indication Ice or snow condition APU fire warning or overheat indication Toilet overservicing Landing gear free-fall condition Overheated wheels Landing-gear tire tread failure
190 06-00 Aircraft areas and dimensions Introduction This chapter describes the aircraft’s general external dimensions, aircraft zoning and station identification. The fuselage stations show the length measurements along the longitudinal axis. All horizontal measurements are taken from the datum line FS 0 which is located at the nose tip. The forward pressure bulkhead is located at station 610 and the fuselage is pressurized between the forward pressure bulkhead and the rear pressure bulkhead, which is located at station 29427.
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Figure 1: Horizontal measurements
FS 0 FS 29427
FS 610
Longitudinal axis
Forward pressure bulkhead
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Rear pressure bulkhead
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190 The aircraft measurement Embraer 190-100 The aircraft has a total length from the nose to the tail of 36.24m, the height from the ground to the top of the vertical tail is 10.55m and a vertical tail area of 16.20 m². The distance from the nose gear to the main gear is 10.60 meters. The aircraft has a total wing span of 28.72m with a total wing area of 92.50 m², a horizontal tail span of 12.01 m, with a horizontal tail area of 26.00 m² and a fuselage external diameter of 3.01 m. The distance from the left to right main gear is 5.94 meters.
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Figure 2: Aircraft measurements Embraer 190 - 100
Total wingspan: 28.72 m
Total wing area: 92.50 m2
Horizontal tail area:26.00 m
2
Horizontal tail span: 12.01 m
10.55 m
Vertical tail area: 16.20 m 2
Distance from nose gear to main gear: 13.83 m
36.24 m
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190 The aircraft measurement Embraer 190 - 200 The aircraft has a total length from the nose to the tail of 38.67m, the height from the ground to the top of the vertical tail is 10.55m and a vertical tail area of 16.20m². The distance from the nose gear to the main gear is 13.83 meters. The aircraft has a total wing span of 28.72 meters with a total wing area of 92.05 m², a horizontal tail span of 12.08 meters, with a horizontal tail area of 26m² and a fuselage external diameter of 3.01 m. The distance from the left to the right main gear is 5.94 meters.
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Figure 3: The aircraft measurement Embraer 190 - 200
13.83 m
38.67 m
12.08 m
5.94 m
28.72 m
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190 The aircraft measurement Located on the left side of the fuselage are two main doors, which qualifies as type 1 emergency exits. Located on the right side of the fuselage are two service doors, which qualify as type 1 exits, and two baggage compartment doors.
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Figure 4: The aircraft measurement
OVERWING EXITS
Type 1 emergency exits
Type 1 exits OVERWING EXITS
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190 The aircraft stations The aircraft stations are defined by a coordinate system using three main reference axes. The point of origin for the longitudinal axis X, laterall axis Y and vertical axis Z is in front of the aircraft. The ordinates are identified by the letter for the major axes, followed by the dimension in inches from the point of origin. There are additional points of origin selected for locating major assemblies. These points are identified with a suffix letter indicating the assembly. These assemblies are the wings, the vertical stabilizer, the horizontal stabilizer, the power plant and the engine pylons.
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Figure 5: Fuselage dimension
EMBRAER 190
FWD Fuselage X=0 X=610
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Center Fuselage 1 X=6232.5
FWD Central Fuselage II
Center Fuselage 2
X=11035 X=13442
AFT Central Fuselage II
Center Fuselage 3
X=19429 X=23146
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Rear Fuselage
X=29632
Tail Cone X=33427 X=36237
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190 Figure 6: Wing station diagram
RIB 1 Y= -1380.00
DRY BAY ENGINE
SLAT 1
RIB 10 YA= -4617.57
WING REF. POINT YA = 00.00
SLAT 2
SLAT 3 WING SPAR 1 SLAT 4 RIB 26 YA= -12196.00
WING SPAR 3 OUTBOARD FLAP
INBOARD FLAP WING SPAR 2
Y = 00.00 YA = 00.00
AILERON WING TRAILING EDGE
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Figure 7: Vertical stabilizer ZV 7020,10
ZV 4269,00
ZV 1069,00
ZV 622,98
Z= 00,00 ZV= 00,00
A
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XV= 00,00
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190
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Figure 8: Horizontal stabilizer Y = 00.00
YH -6100.00
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190 Figure 9: Power plant and engine pylons
X ENG 3877.60
X ENG 5589.90
CANT XP 3256.97 XP 4722.0
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XP 6340.0
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XP 7201.0
XP 8191.0
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The aircraft zoning system The aircraft zoning system provides identification of areas in the aircraft according to 8 major zones, major sub-zones and zones. The major aircraft zones are: • The lower fuselage, zone 100 • The upper fuselage, zone 200 • The tail cone and horizontal and vertical tail, zone 300 • The power plants and pylons, zone 400 • The left wing, zone 500 • The right wing, zone 600 • The landing gear and landing gear wheel well doors, zone 700 • The aircraft doors, zone 800
MAIN LANDING GEAR AND WHELLWELL 700 LOWER FUSELAGE 100 NOSE LANDING GEAR AND WHEELWELL DOOR 700
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POWERPLANT AND PYLONS 400
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Sub zones The major zones are divided into major sub-zones, which are shown by the second digit of the major zone number.The major sub-zones are further divided into zones using the third digit of the major zone number.
Access panels To carry out maintenance on aircraft systems and their components, or to perform inspection of the aircraft structure, adequate access panels and doors are provided in the aircraft surface. Each access panel has an identification number which consists of a three digit zone number followed by two or three letters.The first letter shows the number of the panel within the zone in a logical in a logical sequence and the second letter indicates the location of the panel in relation to the aircraft.An optional third letter is used to identify a floor, wall or ceiling panel. Each panel has a fastener identification code, which identifies the type and the quantity of the fasteners for each panel.
190 7-00 Aircraft jacking Introduction To replace components, the aircraft can be lifted either by using individual landing gear jacking, which permits replacement of wheels and brakes or by completely lifting the aircraft via the jacking points. There are two main jacking points on the wing lower side and one jack point on the rear fuselage lower side. Procedures for lifting a damaged aircraft are described in the “Instructions for Ground Fire Extinguishing and Rescue” manual.
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Figure 1: Landing gear jacking, jacking points
Jacking points NLG
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190 Complete aircraft jacking Complete aircraft jacking, also called three point fuselage jacking, is necessary to perform maintenance such as replacement, repair or functional checks to the landing gear and its components. It can also be used for aircraft weighing.
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Figure 2: Complete aircraft jacking
Complete aircraft jacking (three point fuselage jacking)
190 Before complete aircraft jacking Complete aircraft jacking must be performed in accordance with chapter 7 of the aircraft maintenance manual. Before complete jacking of the aircraft several conditions have to be fulfilled: • All unnecessary equipment below and around the aircraft must be removed. • The aircraft should only be jacked on level ground with the nose pointing into the wind, but preferably in a hangar with closed doors. • Both main landing gears and the nose gear have to be safety locked to prevent inadvertent landing gear retraction. • Install the GSE 070 on all landing gear wheels to prevent the airplane forward and aft movement before the jacks are set. • The emergency/parking brake must be released before jacking. No one is allowed to enter the aircraft during jacking operation. • The aircraft must be lifted to achieve a minimum clearance between the ground and the aircraft tires to assure sufficient clearance for main gear retraction. Note: Before aircraft jacking, refer to the table mentioned in the AMM for weight limitations to ensure that the centre of gravity is within the jacking envelope.
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Figure 3: CG position
CG position
26982
32 54
27900
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190 Jacking points location The aircraft is fully lifted by the use of fuselage jacking points located behind panel 515CB for the LH inboard wing leading edge, and 615CB for the RH inboard wing leading edge. The jacking point for the rear fuselage is located behind panel 313BL. After installing the jack pins and the jacks, the jacks should be operated simultaneously to ensure that jacking is symmetrical and that the aircraft remains level at all time during lifting. When the desired height is reached the jacks have to be locked to prevent inadvertent lowering. To lower the aircraft, the surrounding area has to be cleared of obstructions, the emergency/parking brake must be released and the jacks have to be unlocked. All jacks should then be lowered slowly and symmetrically to ensure that the aircraft maintains a level attitude. As soon as the aircraft weight is off the jacks, wheel chocks should be installed and the emergency/parking brake set.
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Figure 4: Jacking points location
behind panel 515CB for the LH inboard wing leading edge
615CB for the RH inboard wing leading edge
jacking point for the rear fuselage is located behind panel 313BL
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190 Nose and main gear jacking points The nose and main gear jacking points permit individual wheel and brake replacement without weight limitations, and can even be performed with one flat tire at any gear leg. The landing gear safety lock pins have to be put into position, the emergency/parking brake released, and the jack has to be installed below the applicable landing gear leg jacking point. The jack then is raised until the landing gear tire is clear of the ground. After completion of the required maintenance the aircraft can then smoothly be lowered to the ground. When the weight of the aircraft is off the jack it can be removed, the wheels chocked and the emergency/parking brake set.
190 8-00 Leveling and Weighing Introduction Chapter Eight of the Maintenance Manual describes the applicable procedures for leveling and weighing of the aircraft. Only the equipment as specified in the applicable manuals is to be used to perform these tasks. Please note that weighing the aircraft is accomplished using an electronic weighing kit and jacks. Preparation of the aircraft and the weighing procedures are described in the Weight and Balance Manual. Further note that only approved personnel may perform an aircraft weighing.
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Figure 1: Weighing procedures
Chapter Eight of the Maintenance Manual and The Weight and Balance Manual describes the applicable procedures for leveling and weighing of the aircraft
Note that only approved personnel may perform an aircraft weighing
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190 Weighing the aircraft To weigh the aircraft, first install the adapters and the load cells of the electronic weighing kit on the jacks, below the three jacking points on the fuselage. Lift the aircraft until the tires are off the ground, as specified in chapter 07. To determine the weight of the aircraft, refer to the procedures given in the Weight and Balance Manual. After weighing, lower the aircraft as explained in chapter 07 and remove the load cells and the adapters of the electronic weighing kit from the jacks.
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Figure 2: Weighing the aircraft
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190 Aircraft leveling Aircraft leveling is necessary before performing weighing operations, since it permits you to find the accurate centre of gravity of the aircraft. Aircraft leveling is done by lifting the aircraft until the tires are off the ground, as described in chapter 07, and installation of the PLUMB - AIRCRAFT RIGGING kit in the LH main landing gear wheel well.
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Figure 3: Aircraft leveling
A
B
ZONE 1431
WING STUB (REF.)
C
A B
PLUMB BOB CORD ATTACHMENT POINT
GSE039 PLUMB-AIRCRAFT RIGGING (REF.)
AIRCRAFT LEVELING SCALE
C Aircraft Leveling - Maintenance Practices Figure 201
190 9-00 Towing and Taxiing Introduction Chapter nine of the aircraft maintenance manual provides information regarding towing and taxiing of the aircraft. Please note that only approved personnel familiar with the required procedures may tow or taxi the aircraft.
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Figure 1: Chapter 9 Towing and Taxiing
NOTE: Only approved personnel familiar with the required procedures may tow or taxi the aircraft.
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190 Towing the aircraft Towing of the aircraft is performed when it must be moved without using the power of its own engines. To tow the aircraft an approved tow bar with a shear section that breaks at a tension compression shear load is used to prevent damage to the landing gear or the aircraft structure if an excessive load occurs. The aircraft maintenance manual provides a table for the ground towing factors that are most important in various conditions. On this table you can find the necessary drawbar pull and the total wheel-traction load for various aircraft weights, pavement slopes, friction coefficients and engine idle thrust.
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Figure 2: Tow bar and Ground towing requirements table
Tow bar with a shear section
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190 Towing operation Before and during aircraft towing the following safety precautions must be observed. Towing without an approved towbar is prohibited. Alternative devices can cause damage to the aircraft. Before towing: • make sure that down lock safety pins are correctly installed and that all engine cowls are closed and latched. • Disengage the steering system with the switch installed on the control yoke or set the external steering disengagement switch to the "DISENGAGED" position and • make sure that the green towing indication light illuminates. During the towing operation, a technician must stay in the cockpit to set the emergency/parking brake, if necessary. When all towing precautions are preformed which are described in the AMM, you can start the towing operation by the release of the emergency/ parking brake. Tow the aircraft slowly straight ahead before making a turn, and obey the towing speed limitations which are described in the AMM. After completing of the towing operation, tow the aircraft in a straight line for a minimum of 3 meters (10 feet) or until the nose wheel steering system is in its active range of ±76°.
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Figure 3: Before towing
Down lock safety pins installed Engine cowls are closed and latched Disengage the steering system on the control yoke or on the external steering disengagement panel.
Green towing indication light
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190 Aircraft towing Aircraft towing is performed as follows: • Push the lock pin and put the towbar lever in released position. • Install the towbar assembly on the towing attachment on the nose landing gear and push the lock pin and put the towbar lever in towing position. • Attach the other side of the tow bar assembly to the towing vehicle, remove the wheel chocks, and • release the emergency/parking brake. When towing is complete: • • • • •
set the emergency/parking brake, install the wheel chocks and remove the tow bar assembly from the towing vehicle. Push the lock pin and put the towbar lever in released position. Finally, remove the tow bar assembly from the nose landing gear.
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Figure 4: Towbar assembly
TOWING
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190 Towing regulations Only approved persons who are fully familiar with the engine start and shutdown procedures, aircraft limitations, and taxiing techniques are allowed to perform taxiing. In addition, the applicable company procedures and regulations of local authorities must be obeyed. Before aircraft taxiing, clear the area to be used. Install the landing gear safety pins, and make sure that the brakes and the nose wheel steering system are in serviceable condition.
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Figure 5: Towing with towbar
A ZONE 711
TOWING LEVER
LOCKPIN
SHEAR PIN
A
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190 Taxi regulations The following cautions have to be observed: • The areas for taxiing must be free of obstacles and have the necessary space for the manoeuvres. • Always obey the instructions in the operations manual. • Further note that you should not use differential braking during the taxiing. • For the most satisfactory operation, use minimum engine power or, when necessary, slight asymmetric power and the steering control of the nose wheels. The aircraft requires a minimal pavement width for a 180° turn. However, during taxiing you should always perform turns with the largest radius possible given by the available space. Taxi the aircraft at a speed applicable to ramp operations. Also, do not perform turns at a speed greater than 25 kilometers per hour (15 miles per hour). After aircraft taxiing, install the wheel chocks and set the emergency/parking brake.
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Figure 6: Taxi regulations
During taxiing you should always perform turns with the largest radius possible given by the available space.
190 10-00 Parking and Mooring Introduction Chapter 10 of the Maintenance Manual describes the applicable procedures for parking and mooring the aircraft. In general, there are two types of parking: • First, aircraft normal parking, which describes procedures for parking an aircraft for less than 7 days, including parking between flights and overnight. Or, • Aircraft prolonged parking, which describes procedures for parking an aircraft for more than 7 days.
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Figure 1: Types of parking
Two types of parking:
Prolonged parking > 7 days
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190 Minimum distances Please note that the area where the aircraft is parked and moored should be paved and level, with ground tiedown anchors available. Also make sure that there is a minimum distance between the parked aircraft to permit their movement, and that there is a distance of at least 4.5 meters (15 ft) between an operating APU exhaust port and an adjacent aircraft fuel tank vent.
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Figure 2: Minimum distances
Parkingplace must be paved and level, with ground tiedown anchors available.
4.5 m 15 ft
Operating APU exhaust port
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190 For normal aircraft parking For normal aircraft parking make sure that the control handle of the landing gear is in the down position. Make sure that the safety pins are installed on each landing gear. Tow or taxi the aircraft into the position designated for parking. NOTE: Before you park the aircraft, move it in a straight line for approximatel y 3 meters (10 ft), complete the aircraft towing in a straight line for a minimum of 3 m (10 ft) or until the nose wheel steering system is in the range of +/- 76 degrees.. This will remove all torsional stresses applied to the landing gear components and tires during a turn. Ground the aircraft.
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Figure 3: For normal aircraft parking
NOTE: Before you park the aircraft, move it in a straight line for approximately 3 meters (10 ft)
UP
DN LOCK REL
DN
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190 For normal aircraft parking in ice or snow conditions If the parking area has ice or snow on the surface, put one of the following items under the tires: • a mat • a layer of thick sand or other applicable material. This will prevent the tires from freezing to the ground. Set the emergency/parking brake. Retract the flaps if they are extended. Put the chocks against the landing gear wheels and install the covers on the externally mounted aircraft components according to the AMM, using the same procedure as for the air data smart probe covers, the engine inlet covers, the tat sensor covers and the ice detector covers.
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Figure 4: For normal aircraft parking in ice or snow conditions
UP
0 1 B
SLAT/ FLAP
0 1
2
2
3
3
Parking brake set and flaps retracted
R A K E
4
4
5
5
FULL
FULL
DOWN
Ice detector
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TAT sensor
Chapter 10-00
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190 Long term parking Long term parking procedures apply to aircraft that must stay parked for more than seven days. When followed, these procedures will prevent the deterioration of aircraft components exposed to the elements. These components include: • aircraft structure, • airborne equipment/furnishings and • system components. There are different preservation procedures for the different times during which the aircraft must stay out of operation. These times are specified as follows: • Short out-of-operation time - applicable to times from 7 to 60 days and • Long out-of-operation time - applicable to times longer than 60 days. For details regarding these procedures, refer to the appropriate manuals.
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Figure 5: Long term parking
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Long out-of-operation time:
applicable to times from 7 to 60 days
applicable to times longer than 60 days
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190 Mooring the aircraft The procedures for mooring the aircraft are used when the weather conditions are bad or unknown, and/or high wind speeds are expected. For this procedure, tie down rings are installed in each primary brace strut of the main landing gear. Ropes are then used to tie the aircraft to tie down anchors installed in the floor.
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Figure 6: Mooring the aircraft
Tie down rings
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190
Intentionally left blank
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190 ATA 11 Placards and markings
190 Table of Content Placards and Markings 1 Introduction 1
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190
Intentionally left blank
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190 11-00 Placards and Markings Introduction Exterior and interior placards, labels and markings are screen-printed selfadhesive transfer type matte polyester or aluminum metal, attached to the aircraft. Only exterior screen-printed markings, directly on the exterior of the aircraft are protected against contamination by weather, fuel and/or hydraulic fluid by a protective sealer. Some of the labels have there part number printed on them for easier identification and reordering.
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Figure 1: Exterior screen-printed markings
FUEL weather
protective sealer
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fuel
hydraulic fluid
Only exterior screen-printed markings, directly on the exterior of th aircraft are protected against contamination by a protective sealer.
190 12-00 Servicing introduction Chapter 12 of the aircraft maintenance manual provides information about scheduled and unscheduled aircraft servicing, and is divided into the following sub chapters: Replenishing, which provides information about the procedures to fill or charge the aircraft systems with fuel, oil, gas, and other fluids as required. Servicing, which provides information about procedures such as landing gear lubrication and aircraft cleaning. Unscheduled servicing, which provides information about aircraft cold weather maintenance. Please note that you must always refer to the procedures outlined in the applicable manuals to perform these tasks.
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Figure 1: Chapter 12
CHAPTER 12 REPLENISHING
SERVICING
F ill/c h a rg e
P ro c e d u re s
Fuel Oil Gas Other fluids
Landing gear lubrication Aircraft cleaning
UNSCHEDULED
SERVICING
A irc ra ft
Cold weather maintenance
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190 Replenishing Replenishing details the procedures to fill or charge the aircraft systems with fuel, oil, gas, and other fluids as needed. This section contains the following subsections: Fuel tank servicing, Engine and APU servicing, Hydraulic and landing gear system servicing, Oxygen system servicing and water servicing.
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Figure 2: Replenishing
REPLENISHING fill/charge
FUEL
SUBCHAPTERS: FUEL TANK SERVICING
OIL
ENGINE AND APU SERVICING
GAS
HYDRAULIC AND LANDING GEAR SYSTEM SERVICING
OXYGEN SYSTEM SERVICING
OTHER FLUIDS Issue: June06 Revision: 00
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190 Fuel tank servicing Fuel tank servicing can be performed using the following methods: – The fuel tanks can be refuelled/defueled by the use of the pressure refuelling/defueling system, which automatically controls the entire process and stops the refuelling/defueling process at the correct quantity. Pressure refuelling/defueling is performed by the use of the pressure fuelling/defueling adapter and the fuel control panel. – The fuel tanks can be refuelled/defueled by the use of gravity refuelling/defueling. To accomplish this the aircraft has two filler caps on top of the wings. Fuel tank draining for removal of water or other contamination can be carried out by using the drain valves installed at the lowest part of the inboard tanks. The fuel measuring stick assemblies located on each wing lower surface provide a visual indication of the total fuel quantity on each wing. They are used if no electrical power is available to the aircraft or if there is a malfunction of the fuel quantity indicating system. Please note that you must refer to the procedures outlined in the applicable manuals to perform these tasks!
DO NOT INITIATE THE REFUELING BEFORE CONFIRMING THAT THE REFUELING VALVE LIGHTS ARE INITIALLY ON
CLOSED
CLOSED
LH TANK
RH TANK
automatically controls entire process
FUEL QTY REMAINING OPEN
OPEN
OPEN SELECTED
CLOSED
INCR
TKSEL
DECRT
TEST
CLOSED
REFUELING
DEFUELING
GRAVITY
stops at the correct quantity
Pressure fueling/defueling adapter Fuel control panel
MEASURE STICK
Refueling/defueling
Assemblies
Use of gravity
Visual indication
Two filler caps (on top of the wings)
Total fuel quantity
Used if no electr. power is available
FUEL TANK DRAINING Removal of water/ other contamination
if in case of a malfunction of indication system
Drain valves Lowest part ot the inboard tanks
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190 Hydraulic and landing gear system servicing The section "hydraulic and landing gear system servicing" provides information about the servicing procedure on the accumulators of the No. 1, No. 2, and No. 3 hydraulic systems, and also the procedures used to pressurize the landing gear shock struts. Detailed procedures about landing gear servicing can be found in Chapter 32 of the aircraft maintenance manual. Always refer to the procedures outlined in the applicable manuals to perform these tasks!
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Figure 4:
ACCUMULATORS
LANDING GEAR SHOCK STRUT MAIN LANDING GEAR
NOSE LANDING GEAR
Hydr. system #1
Hydr. system #2
Hydr. system #3
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190 Engine and auxiliary power unit servicing The section "engine and auxiliary power unit servicing" provides all required information to refill the engine and auxiliary power unit oil system. Engine oil servicing is performed through the applicable service panels on the engine nacelles, and auxiliary power unit oil servicing is performed through a service panel on the aircraft rear fuselage. Please note that the oil used to service the engine and APU must be listed on the table of approved oils. Also note that you must refer to the procedures outlined in the applicable manuals to perform these tasks.
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Figure 5: Engine and APU servicing
ENGINE OIL SERVICING Service panels Engine nacells
APU OIL SERVICING Service panels Aircraft rear fuselage
! NOTE ! Oil must be listed on the table of approved oils! Issue: June06 Revision: 00
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190 Servicing Water/Waste The section "water/waste" provides information on how to service the water and waste systems. The waste system is serviced through a door installed on the lower right side of the fuselage aft section, while the potable water system is serviced through a door installed on the lower left side of the fuselage aft section. It is recommended that water be removed from the water tank after the last flight of each day if the temperature is expected to fall below freezing. Again, always refer to the procedures outlined in the applicable manuals to perform these tasks!
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Figure 6: Water Waste servicing
NORMAL
DRAIN 1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
MAX FILLING PRESSURE SHALL BE 55 psi 3,8 hpa
FULL
DRAIN
Waste system service door
FILL 1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
1. INSTRUCTION
Potable water system service door
Water must be removed each day after last flight if temperature go below freezing. Issue: June06 Revision: 00
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190 Gaseous servicing The section “gaseous servicing” provides information regarding filling procedures for the cockpit oxygen cylinder, the main and nose gear tire pressure and the hydraulic system accumulator. Attention: Please note that all applicable safety precautions must be obeyed! In addition, you must refer to the procedures outlined in the applicable manuals to perform these tasks!
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Figure 7: Gaseous servicing
Hydr. system #1
Hydr. system #2
Hydr. system #3
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190 Scheduled servicing The section “scheduled servicing” provides information regarding lubricating of the flight controls and landing gear mechanical system, cleaning servicing, like aircraft cleaning, and aircraft disinfect servicing. Refer to the procedures outlined in the applicable manuals to perform these tasks!
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Figure 8: Scheduled servicing
SCHEDULED SERVICING
Flight controls H/S jackscrew
LG mechanical system
A-P servo
Main landing gear
Nose landing gear
Slat track
Main landing gear shockstrut
Cleaning servicing Cleaning-outside
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Chemical cleaning
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190 Unscheduled servicing The section “unscheduled servicing” provides information regarding how to service an aircraft in cold weather conditions or how to perform a cold weather anti-icing and de-icing treatment. Refer to the procedures outlined in the applicable manuals to perform these tasks!
24-64-01 Electrical System Indication Introduction The Electrical Power Generation and Distribution System (EPGDS) communications scheme provides an interface with the Avionics Standard Communications Bus (ASCB), enabling the EPGDS to provide system information via the Modular Avionic Units (MAUs) to the Cockpit Indicating and Crew Alerting system. On the interactive EPGDS communications schematic, you can view the different components, their locations and their functions within the Electrical Power Generation and Distribution System.
Figure 1: Cursor Control Device
Cursor Control Device
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EPGDS Synoptic Page To display the Electrical synoptic page select ELECTRICAL on the MFD using the CCD.
Figure 2: MFD
MFD
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MFD Electrical System Synoptic Page Operation of the electrical system can be monitored by the flight crew on the MFD Electrical System Synoptic Page, which consists of both analog and discrete data. Analog data is used to display actual values like voltage, current and generator frequencies, while discrete data is used to indicate electrical flow paths, conditions of batteries and generators. The CAS field on the EICAS panel informs the crew in case of a system failure by generating a WARNING or CAUTION message and an ADVISORY message to indicate system status. You can now select all warning, caution or advisory messages for display on the EICAS, or you can select the type of information you wish to be displayed on the MFD. By clicking on the messages on the EICAS you will receive additional information regarding their meaning and the system logic behind the message. By clicking on the highlighted symbols on the electrical synoptic page you will receive the related EICAS messages and additional information regarding their meaning.
Figure 3: MFD Electrical Synoptic Page
MFD Electrical Synoptic Page
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MFD Electrical Synoptic Page Engine AC generators • The engine generator icon outline is displayed in green when it is producing power and voltage is above 90 VAC • Normal voltage, frequency and load readouts are green • Generator is white when not producing power Generator has failed, data invalid, or condition is undetermined.
Figure 4: MFD Electrical Synoptic Page
Driven Voltage
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AC External Power - AC GPU is displayed in green if conditions for AC GPU AVAIL have been satisfied. - AC GPU is not displayed if conditions for AC GPU AVAIL have not been satisfied. - AC GPU voltage (V), frequency (Hz), and load (KVA) are displayed in green if conditions for AC GPU AVAIL have been satisfied. - AC GPU voltage (V), frequency (Hz), and load (KVA) are not displayed if conditions for AC GPU AVAIL have not been satisfied.
Figure 5: MFD Electrical Synoptic Page
AC External Power
• External power cart is producing power
• External power cart not producing power
• External power has failed or generator status is undetermined
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EPGDS Synoptic Page Batteries - BATT 1 temperature (oC) is red if BATT 1 temp is greater than or equal to 70 °C (over temp condition). - BATT 1 temperature (oC) is green if BATT 1 temp is less than 70 °C (normal condition). - HOT BATT BUS 1 voltage (V) is always green reflecting values transmitted over ASCB. - BATT 2 temperature (oC) is red if BATT 2 temp is greater than or equal to 70 °C (over temp condition). - BATT 2 temperature (oC) is green if BATT 2 temp is less than 70 °C (normal condition). - HOT BATT BUS 2 voltage (V) is always green reflecting values transmitted over ASCB. - BATT 2 is green if HOT BATT BUS 2 voltage is greater than 18 VDC. - BATT 2 is white if HOT BATT BUS 2 voltage is less than 10 VDC. - BATT 1 is green if HOT BATT BUS 1 voltage is greater than 18 VDC. - BATT 1 is white if HOT BATT BUS 1 voltage is less than 10 VDC.
Figure 6: MFD Electrical Synoptic Page
Batteries • If battery status is invalid the BATT icon is changed to an amber dashed BATT icon • If voltage or temperature is invalid the associated digits change to amber dashes
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EPGDS Communications Architecture
Main Avionics Interface
The Electrical Power Generation and Distribution System (EPGDS) communications scheme provides an interface with the Avionics Standard Communications Bus (ASCB). This interface enables the EPGDS to acquire system information (such as weight on wheels status) from the Modular Avionic Units (MAUs). In addition, it enables the EPGDS to provide maintenance related information to the Central Maintenance Computer (CMC) and crew alert messages to the Engine Indication and Crew Alerting System (EICAS). The GCUs and EPM transmit and receive data as needed for EPGDS coordination via a dedicated 1553 communication bus. An independent Controller Area Network (CAN) communication bus (1553) is provided to allow data to be transferred between the Secondary Power Distribution Assemblies (SPDAs). An ARINC 429 communication scheme is utilized to allow data to be transmitted between the GCUs and SPDAs, and between the SPDAs and Multipurpose Control and Display Units (MCDUs). An additional ARINC 429 bus scheme is implemented to enable communication with the Air Management System (AMS). The GCUs, EPM, and SPDAs have a RS-485 communication interface for their interrogation in a repair shop. Another ARINC 429 interface is utilized to coordinate APU start activity with the APU FADEC. The RS485 communication interface is provided on the GCUs and EPM, and the RS232 communication interface is provided on the SPDAs for interrogation of the controllers during repair shop visit.
The main avionics communications interface used on the ERJ170 is the Avionics Standard Communications Bus (ASCB). The SPDA architecture characterizes ASCB as a high bandwidth interface. As such, the ASCB communication module will connect to the processor modules via the Peripheral Component Interconnect (PCI) bus. Only the active processor will transmit data. Both the active and stand-by processors will read data from the ASCB module. High bandwidth communications protocols such as ASCB cannot be supported adequately on the IEEE-1394 buses. These interfaces are connected to the processors using a 33 MHz, 32 bit, Peripheral Component Interconnect (PCI) bus back plane. The two SPDA processors share the PCI bus segment with the ASCB module.
Cross-Channel Communications The cross-channels communications bus is used to coordinate the utility and load management systems between the SPDAs and as a backup to ASCB. If one SPDA loses its aircraft interface, necessary data will be relayed via the other SPDA to and from the avionics. Only the active processors in each SPDA communicate on the bidirectional cross-channel bus. The cross channel communications between SPDA’s is implemented using the Controller Area Network (CAN) bus operating at 500K bits/second. The cross-channel communications is a dual channel system for redundancy. The total wire length for each of the redundant buses is 262 feet (80 meters) maximum.
Figure 7: Electrical Power Generation and Distribution System (EPGDS) schematic
LICC
RICC
EPM
TX
GCU 2 RX
GCU 1 TX
RX
TX RX
INTER-LRM (1553)
INTERLRM (1553)
MAU 3
AGCU
WIRE SPLICE
TX
RX
ARINC 429 MCDU 1 ARINC 429
TX
RX ARINC 429
PCI
TX
IEE-1394 RX TX
RX
MCDU 2
ARINC 429
TX
RX ARINC 429
RX
RX
TX
TX
SPDA 1
WIRE SPLICE
RX
TX RX RX
RX
ARINC 429 TX RX TX
PCI IEEE-1394
CAN BUS (1553)
RX APU TX FADEC
RX
TX
RS422 SPDA 2
RX TX
AMS
ARINC 429 ASCB
MAU 1
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MAU 2
MAU 3
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CAS Messages List (From FIM 24-00-00) CAS MESSAGE
TYPE
DESCRIPTION
FAULT CODE
AC BUS 1 OFF AC BUS 2 OFF AC ESS BUS OFF
CAUTION
The AC bus 1 is De-energized.
24500100
CAUTION CAUTION
The AC bus 2 is De-energized. The AC bus essential is de-energized.
24500200 24500300
AC STBY BUS OFF
CAUTION
The AC standby bus is de-energized.
24500400
APU GEN OFF BUS
CAUTION
The APU is running, but APU generator is not connected to the AC tie bus. 24200100
BATT 1 DISCHARGING
CAUTION
The battery 1 is discharging.
24300101
BATT 1 OFF
CAUTION
The battery 1 is not connected to DC essential bus 1.
24300301
BATT 1 OVERTEMP
WARNING
The battery 1 has an overtemperature.
24300401
BATT 1-2 OFF
WARNING
The battery 1 and 2 are off.
24300200
BATT 2 DISCHARGING
CAUTION
The battery 2 is discharging.
24300102
BATT 2 OFF
CAUTION
The battery 2 is not connected to DC essential bus 2.
24300302
BATT 2 OVERTEMP
WARNING
The battery 2 has an overtemperature.
24300402
BATT DISCHARGING
WARNING
The battery is discharging.
24300300
BATT 1 TEMP SENS FAULT CAUTION
There is a disagreement between the battery 1 temperature sensors.
24300501
BATT 2 TEMP SENS FAULT CAUTION
There is a disagreement between the battery 2 temperature sensors.
24300502
DC BUS 1 OFF DC BUS 2 OFF DC ESS BUS 1 OFF
CAUTION CAUTION CAUTION
The DC bus 1 is de-energized. The DC bus 2 is de-energized. The DC essential bus 1 is de-energized.
24600100 24600200 24600300
DC ESS BUS 2 OFF DC ESS BUS 3 OFF ELEC EMERGENCY
CAUTION CAUTION WARNING
The DC essential bus 2 is de-energized. The DC essential bus 3 is de-energized. The electrical system is under electrical emergency.
24600400 24600500 24000300
IDG 1 OFF BUS
CAUTION
The IDG 1 is not connected to the AC bus 1.
24200151
IDG 1 OIL
CAUTION
The IDG 1 oil temp is high or its pressure is low.
24200251
IDG 2 OFF BUS IDG 2 OIL INVERTER FAIL LOAD SHED RAT FAIL REMOTE CB TRIP SPDA FAIL TRU 1 FAIL TRU 2 FAIL TRU ESS FAIL
The IDG 2 is not connected to the AC bus 2. The IDG 2 oil temp is high or its pressure is low. The AC inverter is defective. The loas shed was comanded. There is a failure in the RAT system. There is a thermal circuit breaker tripped. There is a defective module in the SPDA. There is a failure in the TRU 1 system. There is a failure in the TRU 2 system There is a failure in the TRU essential system.
Maintenance Message FIM Reference TASK 24-21-00-810-855-A Different GLC 1 Status Sensed by AGCU and GCU 1 A. General (1) This task is for fault code 2421001AGC (GLC1/AGCU/WRG FAULT) (2) The GCU 1 controls the GLC1. B. Fault Description (1) There is a disagreement between the AGCU and the GCU 1 as they sense the GLC 1 status. C. Probable causes (1) Failure of the GLC 1 (AMM MPP 24-21-21/401) (AIPC 24-21-21) (2) Failure of the AGCU (AMM MPP 24-22-03/401) (AIPC 24-22-03) (3) Defective harness (WM 24-61-53) (4) Defective harness in the LICC (AMM MPP 24-51-01/401) (AIPC 24-51-01) (5) Defective harness in the RICC (AMM MPP 24-51-03/401) (AIPC 24-51-03)
Figure 9: MFD CMC Access (Co-Pilot Only)
CCD – CURSOR CONTROL DEVICE
MFD 2
Maintenance System Config
Issue: June06 Revision: 00
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64-
Page 18
170/190 MAINTENANCE TRAINING MANUAL
Electrical System Diagnostic Tests SYSTEM DIAGNOSTICS MENU Select System Diagnostics in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the System Diagnostics Soft Key. • Select the System diagnostics Soft Key by pushing one of the enter keys on CCD no.2 • The System diagnostics menu is displayed and a list of member systems organized by ATA chapter that have system diagnostic pages associated with them are presented.
Figure 10: Electrical System Diagnostic Tests
MAINTENANCE M ESSA GES
SYSTEM DIA GNOSTICS
EXTENDED MAINTENA NCE
DATA LOADER
FILE TRANSFER
Issue: June06 Revision: 00
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64-
Page 20
170/190 MAINTENANCE TRAINING MANUAL
Notes:
Figure 11: Electrical System Diagnostic Tests
Issue: June06 Revision: 00
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64-
Page 22
170/190 MAINTENANCE TRAINING MANUAL
24-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ 42-00 AC External Power C ¦ 1 ¦ System ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦
1) AC GPU AVAIL/IN USE Pushbutton Lights
C ¦ 4 ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
2) AC GPU C ¦ 1 AVAILABLE ¦ Light on ¦ Flight ¦ Attendant ¦ Ground Service ¦ Panel ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
3) AC GPU IN USE C ¦ 1 Light on ¦ Flight ¦ Attendant ¦ Ground Service ¦ Panel ¦
¦ 0 ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
¦ 52-03 In-Flight C ¦ 1 ¦ *** Entertainment ¦ ¦ System (IFE) Auto ¦ ¦ Shutdown ¦
¦ 0 ¦ ¦ ¦
¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 24-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 24 ELECTRICAL POWER ¦ ¦ ¦
¦ 00-05 Batteries 1 and 2 C ¦ 4 ¦ Voltage ¦ ¦ Indication on MFD ¦ ¦ Status Page ¦
¦ 2 ¦ ¦ ¦
¦ One indication per battery may be ¦ inoperative. ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦
C ¦ 4 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ affected battery voltage is ¦ available on MFD Electrical Page.
¦ ¦ ¦
¦ 22-01 APU Generator ¦
C ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided APU ¦ generator remains selected off.
¦ ¦
¦ 36-10 Batteries 1 and 2 C ¦ 4 ¦ Temperature ¦ ¦ Sensors ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 2 ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 41-00 DC External Power D ¦ 1 ¦ *** System ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
1) DC GPU AVAIL/IN USE Pushbutton Lights
D ¦ 2 ¦ ¦ ¦
Any indication on MFD Electrical Page may be inoperative.
| ¦ | ¦ | ¦
One sensor per battery may be inoperative provided at least one temperature of associated battery on Electrical Synoptic Display (MFD Electrical Page) is verified to operate normally before each flight.
(O)May be inoperative provided IFE RACK Power Switch is verified to operate normally before each departure.
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
RICC
AC BUS 2 CB1 CB29 35A 50A
GSTC 60A
AC GND SVC
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 1
Figure 2: Batteries only RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A
TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2
AC INVERTER 250VA DC
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
ABC 400A
HOTBATT BUS 1
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 2
170/190 MAINTENANCE TRAINING MANUAL
Figure 3: Battery 2 supplying APU start function RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A
TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
DC ESS BUS 3
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 3
Figure 4: DC EXT power supplying APU function RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A
STANDBY AC BUS TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
EF1 150A
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
DC ESS BUS 3
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 4
170/190 MAINTENANCE TRAINING MANUAL
Figure 5: APU generator supplying A/C RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
GSTC 60A
50A
AC GND SVC TRU ESS 300A
EF2 150A
ETC1 120 A
50A
TRU2 300A
TRU 1C 400A DC BUS 1
DC ESS BUS 3
CB1
35A
35A TRU 1 300A
EC 1 120A
EF1 150A
CB29
CB4
DC GND SVC
TRU EC 400A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A CB17
EF4 225A
15A
AF1 225A
AICC
HOTBATT BUS 2
AC INVERTER 250VA DC VA DC 250
FOR TRAINING ONLY Reproduction Prohibited
BATT 11
ISOLATED RICC AF2 ASC 300A 400A
ABC 400A
HOTBATT BUS 1
Issue: June06 Revision: 01
RICC
AC BUS 2
CB3
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25 A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
EPDC 400A
TO APU START DC EXT PWR
BATT 22
Chapter 24-64
Page 5
170/190 MAINTENANCE TRAINING MANUAL
Figure 6: APU GEN Supplying Airplane ALC Close Command
APU GEN Switch ON Protective trip of APU Gen channel AND Tie Bus dead APU Gen power ready conditions satisfied AGCU granted tie bus access
ALC Close Command
OR AND
Power transfer control logic (autoparallel) AGCU
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64
Page 6
170/190 MAINTENANCE TRAINING MANUAL
Figure 7: A/C EXT power supplying aircraft RAT GEN
GLC 1 120A
EICC RLC 60A
EPAC 120A
STANDBY AC BUS TRU ESS 300A
DC ESS BUS 3
EF2 150A
CB4 35A TRU 1 300A
TRU2 300A
TRU 1C 400A DC BUS 1
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
EC 1 120A
EF1 150A
GLC 2 120A
BTC 2 120A
DC GND SVC
TRU EC 400A
IDG 2
ALC 120A
LICC
AC BUS 1 CB3 50A AC GND SVC
CB26 25A
STBYC 10A
APU GEN
BTC 1 120A
AETC 60A
AC ESS BUS
IN USE
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 7
Figure 8: AC EXT PWR Supplying Airplane BTC Close Command _ Primary control
AC BUS TIES Switch AUTO Associated BTC trip command Associated BTC inhibit signal Associated AC Main Bus is dead Another power source available to supply associated AC Main Bus
AND AND
BTC Close Command
Associated GCU allows BTC1 closure Associated GCU is granted tie bus access Power transfer control logic (autoparallel) Is commanding associated BTC to close
AND
OR
AND
Associated GCU is granted tie bus access GCU1 or GCU2
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64
Page 8
170/190 MAINTENANCE TRAINING MANUAL
Figure 9: AC EXT PWR Supplying Airplane BTC Close Command- Secondary control
GCU1/2 loss of power GCU1/2 failsafe
OR
Hardware discrete inhibit from GCU1/2 1553 software discrete inhibit from GCU1/2 AC BUS TIES Switch AUTO
AND
Associated BTC trip command
BTC Close Command
Associated BTC inhibit signal Associated AC Main Bus is dead Another power source available to supply associated AC Main Bus
AND OR
AGCU allows BTC1 closure AGCU is granted tie bus access
AND AGCU
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64
Page 9
170/190 MAINTENANCE TRAINING MANUAL
Figure 10: A/C EXT power supplying ground SVC buses
RAT GEN
GLC 1 120A
EICC RLC 60A
AVAIL
APU GEN
EPAC 120A
AC BUS 1
25A
TRU ESS 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
CB29 35A
GSTC 60A
CB4 35A TRU 1 300A
DC BUS 1
ETC1 120 A
CB17 15A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC
BC2 200A
DC ESS BUS 1
BC 1 200A
CB1 50A
TRU2 300A
TRU 1C 400A EC 1 120A
EF1 150A
RICC
AC BUS 2
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
CB3 50A
CB26
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 10
170/190 MAINTENANCE TRAINING MANUAL
Figure 11: IDGs supplying aircraft (Ground Mode)
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
AC BUS 1
TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
CB3 50A
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 11
Figure 12: IDGs Supplying Airplane (ground mode) GLC Close Command
GEN CONTROL Switch ON Protective trip of associated IDG channel Associated IDG power ready conditions satisfied Associated AC Main Bus dead Tie Bus dead No other power source available to power tie bus
AND
OR
AND
GLC Close Command
Power transfer control logic (autoparallel)
Associated BTC commanded open OR Associated GCU granted tie bus access GCU1 or GCU2 Issue: June06 Revision: 01
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 14
170/190 MAINTENANCE TRAINING MANUAL
Figure 15: IDGs Supplying Airplane (air mode) Air Mode Determination
SPDA1 EEC1
ARINC 429
Weight On Wheels (WOW) = Air Mode Engine 1 speed > 12,750 RPM (75% N2)
OR
1 sec
Air Mode GCU1
Inter-LRM (1553)
SPDA2 EEC2
Weight On Wheels (WOW) = Air Mode Engine 2 speed > 12,750 RPM (75% N2)
OR
1 sec
Air Mode GCU2
Inter-LRM (1553) ASCB Weight On Wheels (WOW) = Air Mode Engine 2 speed > 12,750 RPM (75% N2) SPDA2
OR
1 sec
Air Mode
Engine 1 speed > 12,750 RPM (75% N2) AGCU
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64
Page 15
Figure 16: IDG 1 and APU GEN supplying aircraft
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
CB4 35A TRU 1 300A
TRU2 300A
TRU 1C 400A DC BUS 1
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
EC 1 120A
EF1 150A
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 16
170/190 MAINTENANCE TRAINING MANUAL
Figure 17: IDG 1 supplying aircraft
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A
TRU ESS 300A
3PH 35A TRU 1 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
EF3 200A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
BC2 200A
DC ESS BUS 1
BC 1 200A
CB17 15A
EF4 225A
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 17
Figure 18: IDG 1 Supplying Airplane BTC Close Command
AC BUS TIES Switch AUTO Associated BTC trip command Associated BTC inhibit signal Associated AC Main Bus is dead Another power source available to supply associated AC Main Bus
AND AND
BTC Close Command
Associated GCU allows BTC1 closure Associated GCU is granted tie bus access Power transfer control logic (autoparallel) Is commanding associated BTC to close
AND
OR
AND
Associated GCU is granted tie bus access GCU1 or GCU2
Issue: June06 Revision: 01
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-64
Page 18
170/190 MAINTENANCE TRAINING MANUAL
Figure 19: RAT and batteries supplying the airplane (Air Mode) RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
AC BUS 1
TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
DC ESS BUS 3
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
CB17 15A
RF1 150A
DCTC 120A ETC 2 120A
DC ESS BUS 2
EF3 200A DC ESS BUS 1 BC 1 200A EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A BC2 200A AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 01
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
STANDBY AC BUS
GLC 2 120A
BTC 2 120A
CB3 50A
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 19
Figure 20: IDGs supplying airplane (Air Mode), TRU 2 failed
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB26 25 A
STANDBY AC BUS TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
Failure
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
DC ESS BUS 1 BC 1 200A EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A BC2 200A AF1 225A
AICC
HOTBATT BUS 2
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BATT 1
ISOLATED RICC AF2 ASC 300A 400A
ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
TRU2 300A
DC ESS BUS 2
EF3 200A
CB17 15A
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
GLC 2 120A
BTC 2 120A
AC BUS 1
CB3 50A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
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170/190 MAINTENANCE TRAINING MANUAL
Figure 21: TRU1C or TRU2C Close Command Primary control
TRU Switch latched in (ON) AC Main Bus powered TRU overcurrent condition
AND
TRU1C or TRU2C Close Command
TRU output current no longer present TRU voltage ripple excessive GCU1 or GCU2
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Chapter 24-64
Page 21
Figure 22:TRU1C or TRU2C Close Command Secondary control
GCU (1 or 2) loss of power GCU (1 or 2) failsafe
OR
Hardware discrete inhibit from GCU (1 or 2) 1553 software discrete inhibit from GCU (1 or 2) AND
TRU Switch (1 or 2) ON
TRU (1C or 2C) Close Command
AC Main Bus (1 or 2) powered TRU (1 or 2) overcurrent condition TRU (1 or 2) output current no longer present TRU (1 or 2) voltage ripple excessive AGCU
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170/190 MAINTENANCE TRAINING MANUAL
Figure 23: IDGs supplying airplane (Air Mode), AC bus 1 short
RAT GEN
AC EXT PWR
IDG 1
3 -trip
GLC 1 120A
EICC RLC 60A
CB3 50A
TRU ESS 300A
DC GND SVC
TRU EC 400A
DC ESS BUS 3
EF2 150A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
EF1 150A
ETC1 120 A
RF1 150A
DCTC 120A ETC 2 120A
DC ESS BUS 1 BC 1 200A EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A DC ESS BUS 2
EF3 200A
CB17 15A
BC2 200A AF1 225A
AICC
HOTBATT BUS 2
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BATT 1
ISOLATED RICC AF2 ASC 300A 400A
ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
RICC
AC BUS 2 CB29 CB1 35A 50A
CB4 35A TRU 1 300A
1 - short to ground
GLC 2 120A
BTC 2 120A
GSTC 60A
AC GND SVC
IDG 2
ALC 120A
LICC
AC BUS 1
CB26 25A
STANDBY AC BUS
EPAC 120A
APU GEN
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
2 - lock out
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 23
Figure 24: IDGs supplying airplane (Air Mode), DC bus 2 short
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
AC BUS 1
AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A TRU 1C 400A
EF1 150A
EF2 150A
CB17 15A
EF3 200A ESS BUS 1 BC 1 200A EF4 225A
TRU2 300A
3 - Lockout
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A RF1 150A
DC BUS 1
ETC1 120 A
DCTC 120A
DC ESS BUS 2
2 - Blows open
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BATT 1
FOR TRAINING ONLY Reproduction Prohibited
BC2 200A AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
EC 1 120A DC ESS BUS 3
GLC 2 120A
BTC 2 120A
CB3 50A
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
4 - Open 1 - Short to ground 4 - Open ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
Page 24
170/190 MAINTENANCE TRAINING MANUAL
Figure 25: IDGs supplying airplane (Air Mode), DC ESS bus 3 short RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
AC BUS 1
STANDBY AC BUS TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
TRU2 300A
TRU 1C 400A
RF1 150A
DC BUS 1 EC 1 120A
DC ESS BUS 3
EF1 150A
ETC1 120 A
EF2 150A
3 - Open
DCTC 120A ETC 2 120A
DC ESS BUS 2
EF3 200A ESS BUS 1 BC 1 200A
CB17 15A
EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 200A BC2 200A
3 - Open
AF1 225A
AICC
HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
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RICC
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
AC GND SVC
2 - Trips
GLC 2 120A
BTC 2 120A
CB3 50A
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
FOR TRAINING ONLY Reproduction Prohibited
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-64
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Intentionally left blank
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170/190 MAINTENANCE TRAINING MANUAL
24-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ 42-00 AC External Power C ¦ 1 ¦ System ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦
1) AC GPU AVAIL/IN USE Pushbutton Lights
C ¦ 4 ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
2) AC GPU C ¦ 1 AVAILABLE ¦ Light on ¦ Flight ¦ Attendant ¦ Ground Service ¦ Panel ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
3) AC GPU IN USE C ¦ 1 Light on ¦ Flight ¦ Attendant ¦ Ground Service ¦ Panel ¦
¦ 0 ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
¦ 52-03 In-Flight C ¦ 1 ¦ *** Entertainment ¦ ¦ System (IFE) Auto ¦ ¦ Shutdown ¦
¦ 0 ¦ ¦ ¦
¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 24-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 24 ELECTRICAL POWER ¦ ¦ ¦
¦ 00-05 Batteries 1 and 2 C ¦ 4 ¦ Voltage ¦ ¦ Indication on MFD ¦ ¦ Status Page ¦
¦ 2 ¦ ¦ ¦
¦ One indication per battery may be ¦ inoperative. ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦
C ¦ 4 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ affected battery voltage is ¦ available on MFD Electrical Page.
¦ ¦ ¦
¦ 22-01 APU Generator ¦
C ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided APU ¦ generator remains selected off.
¦ ¦
¦ 36-10 Batteries 1 and 2 C ¦ 4 ¦ Temperature ¦ ¦ Sensors ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 2 ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 41-00 DC External Power D ¦ 1 ¦ *** System ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
1) DC GPU AVAIL/IN USE Pushbutton Lights
D ¦ 2 ¦ ¦ ¦
Any indication on MFD Electrical Page may be inoperative.
| ¦ | ¦ | ¦
One sensor per battery may be inoperative provided at least one temperature of associated battery on Electrical Synoptic Display (MFD Electrical Page) is verified to operate normally before each flight.
(O)May be inoperative provided IFE RACK Power Switch is verified to operate normally before each departure.
24-00 Electrical power general Introduction The Electrical Power Generating and Distribution System (EPGDS) is comprised of AC and DC power sources. The AC system consists of • • • •
two engine-driven Integrated Drive AC Generators (IDGs), one Auxiliary Power Unit driven AC-Generator (APU GEN); one Ram Air Turbine driven AC-Generator (RAT) and one AC external power input.
The DC system consists of: • one DC external power input and • two nickel cadmium accumulators. Normal operation of the EPGDS is in automatic mode, whereby selection of the power source for each bus is accomplished automatically.
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Chapter 24-00
Page 1
Figure 1: Electrical Power Generating and Distribution System (EPGDS) schematic
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
3PH 35A
AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
3PH 35A TRU 1 300A
TRU2 300A
TRU 1C 400A DC BUS 1
ETC1 120 A
1PH 15A
DCTC 120A
EF3 200A ESS BUS 1 BC 1 200A EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A ESS BUS 2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
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RICC
AC BUS 2 3PH 3PH 50A 35A
GSTC 60A
EC 1 120A ESS BUS 3 EF1 EF2 150A 150A
GLC 2 120A
BTC 2 120A
AC BUS 1 3PH 50A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
FOR TRAINING ONLY Reproduction Prohibited
ISOLATED AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-00
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170/190 MAINTENANCE TRAINING MANUAL
EPGDS Component Locations Each IDG is installed on its respective engine gearbox. The APU Generator is installed on the APU gearbox. The following EPGDS equipment is installed in the pressurized forward avionics bay: - Essential Integrated Control Centre (EICC)
- Auxiliary Generator Control Unit (AGCU) - Transformer Rectifier Unit # 2 (TRU 2) - Secondary Power Distribution Assembly # 2 (SPDA 2) The following EPGDS equipment is installed in the pressurized rear avionics bay:
- Transformer Rectifier Unit Essential (TRU ESS)
- Auxiliary Integrated Control Centre (AICC)
- Secondary Power Distribution Assembly # 1 (SPDA 1)
- Battery # 2 (BATT 2)
- Battery #1 (BATT1) - Static Inverter The following EPGDS equipment is installed in the pressurized mid avionics bay:
The Ram Air Turbine (RAT) system is installed adjacent to the nose wheel bay on the right side of the aircraft outside of the forward Ebay just below the co-pilot position. The following LRUs are accessible from the nose wheel bay: - RAT Actuator
- Left Integrated Control Centre (LICC) - Restow Pump - Generator Control Unit # 1 (GCU 1) - Uplock - External Power Module (EPM) - Manual Release Cable Assembly - Transformer Rectifier Unit # 1 (TRU 1) The RAT GCU is installed in the forward pressurized avionics bay. - Right Integrated Control Centre (RICC) - Generator control Unit # 2 (GCU 2)
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Chapter 24-00
Page 3
Figure 2: Electrical system
LICC
IDG 1
TR U
Static Inverter
SPDA 1
SPDA 2 EICC
APU Generator
Battery 1
Battery 2
AICC
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IDG 2
RAT System
Chapter 24-00
Page 4
170/190 MAINTENANCE TRAINING MANUAL
Integrated Control Centres (ICCs) The Left and Right Integrated Control Centres are located in the temperature and pressure controlled Mid E-Bay. Access to this equipment can is through the Mid E-Bay floor access hatch, located on the aircraft port side behind the left wing. Both ICCs provide control, protection and distribution of primary AC / DC electrical power.
The ICC cooling The ICCs are mounted and bonded using four bolts through the four mounting feet. Cooling air is provided from the aircraft Air Management System (AMS) and drawn through the TRU to ensure adequate heat dissipation. All other ICC components are cooled by natural convection. All normal maintenance for removal and replacement of the TRUs, GCUs, EPM, contractors, relays, fuses and circuit breakers is performed through the front of both ICCs.
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Chapter 24-00
Page 5
Figure 3: LICC and RICC Locations
Both ICCs provide ntrol tection distribution mary AC / DC electrical power
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Mid E-Bay floor access hatch
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-00
Page 6
170/190 MAINTENANCE TRAINING MANUAL
LICC (GCU 1, EPM, and TRU 1) The ICCs share some common design features, including the use of Line Replaceable Units (LRMs). The same Generator Control Unit (GCU) is used in the LICC and RICC. There is a common Transformer Rectifier Unit (TRU) in the LICC, RICC, and EICC. The contactor base plate assembly is similar in the three ICCs. Common circuit breakers, which are LRMs, are used in all ICCs. Circuit breakers are mounted and accessible through the front panel of the ICCs. Relays are common and are socket mounted LRMs. The LICC provides control, protection, and distribution of primary AC/DC power. Scheduled maintenance of the LICC is not required. The LICC contains the following LRMs accessible from the front face without opening circuit breaker panels: GCU 1: Generator Control Unit # 1 provides control, protection and distribution of power generated by IDG 1. EPM: External Power Module provides protection and distribution of external AC power. TRU 1: Transformer Rectifier Unit # 1 provides DC power to DC BUS 1, DC ESS BUS 1, DC GND SVC Bus, backup power for DC BUS 2 and DC ESS BUS 3, and charging current for Battery 1 under normal configuration operation.
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Page 7
Figure 4: LICC
TRU 1
GCU 1
EPM
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Page 8
170/190 MAINTENANCE TRAINING MANUAL
LICC (Circuit Breakers and Fuses) The LICC provides AC and DC power distribution and protection for the main aircraft busses associated with the left side of the aircraft. The circuit breakers are thermal type devices with temperature compensation. When subjected to an overload current, the breaker will trip open after a predetermined time. The circuit breakers are the free tripping, push/pull, on/off, manual actuation type. The CB aux contacts are normally open polarized with a blocking diode. These CBs are accessible by opening the associated front panels. Removal and replacement can be accomplished by loosening the attaching hardware and the interface wire harness. Precautions should be taken to ensure proper maintenance for power considerations are practised while conducting maintenance on the circuit. These devices are not to be resettable or replaced in flight. The fuses are thermal type devices, which melt when subjected to an overload current. The fuse blown detector is mounted in parallel to the fuse. When the fuse melts, the fuse blown detector changes permanently the position of a “form C” contact. The fuse and fuse blown detector can only be used one.
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Chapter 24-00
Page 9
Figure 5: LICC Internal
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170/190 MAINTENANCE TRAINING MANUAL
RICC (GCU 2, AGCU, and TRU 2) The ICCs share some common design features, including the use of Line Replaceable Units (LRMs). The same Generator Control Unit (GCU) is used in the LICC and RICC. There is a common Transformer Rectifier Unit (TRU) in the LICC, RICC, and EICC. The contactor base plate assembly is similar in the three ICCs. Common circuit breakers, which are LRMs, are used in all ICCs. Circuit breakers are mounted and accessible through the front panel of the ICCs. Relays are common and are socket mounted LRMs. The RICC provides control, protection, and distribution of primary AC/DC power. Scheduled maintenance of the RICC is not required. The RICC contains the following LRMs accessible from the front face without opening circuit breaker panels: GCU 2: Generator Control Unit # 2 provides control, protection, and distribution of power generated by IDG2. AGCU: Auxiliary Generator Control Unit provides control, protection, and distribution of power generated by the APU Generator. TRU 2: Transformer Rectifier Unit # 2 provides DC power to DC BUS 2, DC ESS BUS 2, backup power for DC BUS 1 and DC ESS BUS 3, and charging current for Battery 2 under normal configuration operation.
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Page 11
Figure 6: RICC
GCU 2
TRU 2
AGCU
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Page 12
170/190 MAINTENANCE TRAINING MANUAL
RICC (Circuit Breakers and Fuse) CAUTION: All power sources to the ICC need to be disconnected prior to performing any maintenance on the ICC (such as opening the panels or removing any LRMs). The circuit breakers are thermal type devices with temperature compensation. When subjected to an overload current, the breaker will trip open after a predetermined time. The circuit breakers are the free tripping, push/pull, on/off, manual actuation type. The CB aux contacts are normally open polarized with a blocking diode. These CBs are accessible by opening the associated front panels. Removal and replacement can be accomplished by loosening the attaching hardware and the interface wire harness. Precautions should be taken to ensure proper maintenance for power considerations are practised while conducting maintenance on the circuit. These devices are not to be resettable or replaced in flight. The fuse is a thermal type devices, which melts when subjected to an overload current. The fuse blown detector is mounted in parallel to the fuse. When the fuse melts, the fuse blown detector changes permanently the position of a “form C” contact. The fuse and fuse blown detector can only be used one.
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Chapter 24-00
Page 13
Figure 7: RICC Internal
DC Contactor & Relay Cavity AC Contactor Cavity
AC Relays
DC Circuit Breakers
AC Circuit Breakers
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170/190 MAINTENANCE TRAINING MANUAL
The Essential Integrated Control Centre (EICC) The Essential Integrated Control Centre (EICC) is located in the temperature and pressure controlled Forward E-Bay. Access to this equipment is through the forward E-Bay floor access hatch, which is located in front of the nose gear. The EICC provides control, protection and distribution of primary AC/ DC electrical power. The EICCs are mounted and bonded using four bolts trough the four mounting feet. Cooling air is provided by the Air Management System (AMS) and drawn through the TRU to ensure adequate heat dissipation. All other EICC components are cooled through natural convection. All normal maintenance for removal and replacement of the Essential TRU (TRU ESS), contractors, relays, fuses and circuit breakers is performed through the front of the EICC. Precautions should be taken to ensure proper power safety procedures are practised while conducting maintenance on the EICC.
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Chapter 24-00
Page 15
Figure 8: EICC
Forward E-bay
EICC EICC
EICC provides trol, tection distribution ry AC / DC electrical power.
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170/190 MAINTENANCE TRAINING MANUAL
EICC (Circuit Breakers)
EICC (TRU ESS)
All loads is sourced from AC ESS BUS, DC ESS BUS 1, DC ESS BUS 3, HOT BATT BUS 1, or STANDBY AC BUS are protected by a circuit breaker on the EICC. The status of these circuit breakers is monitored by SPDA 1 and annunciated in the cockpit by means of a CAS message (REMOTE CB TRIP) and by means of the Multi functional Functional Display Unit (MCDU). The circuit breakers are thermal type devices with temperature compensation. When subjected to an overload current, the breaker will trip open after a predetermined time. The circuit breakers are the free tripping, pull/push, on/off, manual actuation type. The CB aux contacts are normally open polarized with a blocking diode. These CBs are accessible by opening the associated front panels. Removal and replacement can be accomplished by loosening the attaching hardware and the interface wire harness. Precautions should be taken to ensure proper maintenance for power considerations are practised while conducting maintenance on the circuit. These devices are not to be resettable or replaced in flight. CAUTION: All power sources to the ICC need to be disconnected prior to performing any maintenance on the ICC (such as opening the panels or removing any LRMs)
The ICCs share some common design features, including the use of Line Replaceable Units (LRMs). There is a common Transformer Rectifier Unit (TRU) in the LICC, RICC and EICC. The contactor base plate assembly is similar in the three ICCs. Common circuit breakers, which are LRMs, are used in all ICCs. Circuit breakers are mounted and accessible through the front panel of the ICCs. Relays are common and are socket mounted LRMs. The EICC provides control, protection, and distribution of primary AC/DC power. Scheduled maintenance of the EICC is not required. The following LRM is accessible from the front face of the EICC without opening circuit breaker panels: TRU ESS: The Essential Transformer Rectifier Unit provides DC power to DC ESS BUS 3. The TRU ESS can be replaced after removing the 4 captive screws. CAUTION: All power sources to the ICC need to be disconnected prior to performing any maintenance on the ICC (such as opening the panels or removing any LRMs).
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Page 17
Figure 9: EICC Internals
DC Circuit Breakers
DC Contactor & Relay Cavity
TRU ESS Cavity
AC Monitor Sensor K4 Cavity
AC Circuit Breakers
AC Contactor & Relay Cavity
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170/190 MAINTENANCE TRAINING MANUAL
The Auxiliary Integrated Control centre (AICC) The Auxiliary Integrated Control Centre (AICC) is located in the temperature and pressure controlled Aft E-Bay. Access to this equipment is through an airplane panel located in the rear of the cabin. The AICC provides control, protection and distribution of primary AC/DC electrical power. The AICCs are mounted and bonded using four fasteners through the four mounting feet. All AICC components are cooled using natural convection. All normal maintenance for removal and replacement of the contactors, relays, fuses and circuit breakers is performed through the front of the AICC. Precautions should be taken to ensure proper power safety procedures are practised while conducting maintenance on the AICC.
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Page 19
Figure 10: AICC
AFT E-Bay
AICC Battery 2
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170/190 MAINTENANCE TRAINING MANUAL
AICC Electrical Components The AICC contains the following LRMs depicted in the electrical schematic: EPDC: The External Power DC Contactor provides an interface between DC EXT PWR and the APU START BUS. It is controlled automatically through DC external power relay logic. The EPDC closes when the following conditions exist: - Acceptable external DC power quality exists - DC GPU PWR switch AVAIL lamp illuminates - DC GPU PWR switch is latched ON The EPDC opens when the DC GPU PWR switch is unlatched OFF. ASC: The APU Start Contactor allows BATT 2 or DC EXT PWR to be routed to the APU for starting. It is controlled by SPDA 2. ABC: The APU Start Bus Contactor allows BATT 2 power to be routed to the APU START BUS. It is controlled by SPDA 2. The HOT BATT BUS 2 provides a point of distribution for BATT 2 power. The APU Start Bus provides an interface to the APU and is powered by DC EXT PWR or BATT 2 during the APU start cycle.
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Figure 11: AICC Components
APU Start Contactor (ASC) Fuse (AF2)
External Power DC Contactor (EPDC)
Circuit Breakers (15)
Voltage Sense Relay (K1 )
APU Bus Contactor (ABC) Issue: June06 Revision: 00
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170/190 MAINTENANCE TRAINING MANUAL
24-20-00 AC Generation general Introduction The AC system of the airplane operates on 115 Volts AC, at 400Hz frequency stabilized. There are three main AC power sources: • two Integrated Drive Generators (IDGs), installed on their respective engine gearboxes, and one APU Generator installed on the APU gearbox. • There is also an AC external power source interface for ground operation, • a Ram Air Turbine (RAT) for emergency purposes in flight, and • a Static Inverter which goes into operation when only one main AC power source is available.
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Chapter 24-20-
Page 1
Figure 1: AC Power Generation
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170/190 MAINTENANCE TRAINING MANUAL
AC BUS TIES The AC BUS TIES switch provides control of Bus Tie Contactor 1/2 (BTC 1 and BTC 2). With the switch in the AUTO position, the EPGDS will operate the BTC 1 and BTC 2 automatically. With the switch in the 1 OPEN position, BTC 1 will open and BTC 2 will operate automatically. With the switch in the 2 OPEN position, BTC 1 will operate automatically. For some system protections a BTC1 and / or BTC2 lockout system is operative. To reset a lockout, the AC BUS TIES switch must be rotated from the AUTO position to the respective BTC position and back to AUTO position.
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Figure 2: AC Bus tie contactors
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A BTC 1 120A
AETC 60A
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
DC ESS BUS3 EF1 EF2 150A 150A
DC BUS 1
ETC1 120 A
CB17 15A
EF3 200A DC ESS BUS 1 BC 1 200A EF4 225A
DCTC 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
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BATT 1
operate automatically
TRU2 300A
TRU 1C 400A EC 1 120A
RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
AC GND SVC
GLC 2 120A
BTC 2 120A
open
CB3 50A
CB26 25A
IDG 2
ALC 120A
LICC
AC BUS 1
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
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NOTES:
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Figure 3: AC Bus Tie Contactors
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A BTC 1 120A
AETC 60A
STANDBY AC BUS TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
DC ESS BUS3 EF1 EF2 150A 150A
DC BUS 1
ETC1 120 A
EF3 200A DC ESS BUS 1 BC 1 200A
CB17 15A
EF4 225A
DCTC 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
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BATT 1
open
TRU2 300A
TRU 1C 400A EC 1 120A
RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
AC GND SVC
GLC 2 120A
BTC 2 120A
operate automatically
CB3 50A
CB26 25A
IDG 2
ALC 120A
LICC
AC BUS 1
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
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Load Shed The load shed/restoration function (logic) is automatically implemented and controlled in the SPDA(s) utilizing information sent over the ARINC 429 and ASCB communication buses to identify and initiate the shed/restoration function. The SPDA utilizes the following EPGS status information: - ARINC 429 information from the GCUs and EPM - To determine system configuration (number of AC power sources on-line based on contactor information) - ASCB information from the MAUs - To determine Weight On Wheels status (air/ground mode) - ARINC 429 information - To determine individual generator load information When a single generator (IDG or APU GEN) is operating while the aircraft is in air mode, the load shed function shall simultaneously shed all the non-essential loads as shown. Additionally, independent of air/ground mode configuration, the SPDA(s) shall interpret load information being provided by the associated GCUs. If generator phase current load information indicates: 116A = generator single phase current < 130A for 2.5 minutes OR 130A = generator single phase current < 174 A for 2.5 seconds then the load shed function shall simultaneously shed all the nonessential loads as shown.
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Figure 4: Single Generator Operation
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NBPT - No break Power Transfer COORDINATION OF POWER TRANSFERS ON GROUND SYSTEM ALLOWS TWO AC SOURCES TO BE CONNECTED TO THE SAME BUS FOR A LIMITED PERIOD DURING POWER SOURCE CHANGE OVER.
associated contactor without any time delay depending on whether a power transfer is feasible at the time. The time allowed for IDG related NBPTs is 5 seconds. The time allowed for NBPTs between the APU GEN and an IDG is 15 seconds. The time allowed for NBPTs between the AC EXT PWR source and an IDG or the APU GEN is 15 seconds. IF an NBPT does not occur within the allowed time delay, the transfer shall be accomplished by means of a BPT with a minimum power interruption.
TO PREVENT SYSTEM INTERRUPTION (EG BLANKING SCREENS) - IDG RELATED NBPT - TIME ALLOWED - 5 SECS - APU GEN OR EXT POWER NBPT - TIME ALLOWED - 15 SECS. During normal operation, the two IDGs operate as the primary AC power sources in a split bus configuration supplying their respective AC BUSes. IGD 1 supplies AC BUS 1 when the Generator Line Contactor #1 (GLC1) is commanded closed by Generator Control Unit#1 (GCU1). IDG2 supplies AC BUS 2 when GLC2 is commanded closed by GCU2. The AC Tie Bus connects AC BUS 1 and AC BUS 2 enabling the opposite IDG, the APU GEN, or an AC EXT PWR source to supply an AC BUS. It is powered when AC EXT PWR is connected to the aircraft, and the External Power Contactor (EPAC) is commanded closed by the External Power Module (EPM). It is also powered during No Break Power Transfers (NBPTs), which occur during APU and engine start on the ground. During flight it is powered only when an IDG fails or if an engine is shut down. The APU GEN can deliver sufficient power to serve as an alternate source of power in flight if one or both IDGs fail. The EPGDS is designed to coordinate NBPTs on the ground between IDGs, AC EXT PWR, and the APU GEN. Power transfers in the air are break power transfers (BPTs). Positioning the cockpit control panel switches to the auto position enables the associated AC power source for transfer of power according to availability and bus priority rules. Manually selecting a system control switch to the OFF position either initiates a power transfer or trips the Issue: June06 Revision: 00
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Figure 5: NBPT No Break Power Transfer AC EXTERNAL POWER
IDG 1 MIDDLE AVIONICS COMPT
RAT GEN
APU GEN
(SSM 24-42-80)
(SDS 24-21) (MPP 24-21-01)
GCU 1
EPM GLC 1
(SDS 24-23) (MPP 24-23-01)
RAT GCU
FORWARD AVIONICS COMPT
BTC 1
IDG 2 (SDS 24-22) (MPP 24-22-01)
MIDDLE AVIONICS COMPT
(SDS 24-21) (MPP 24-21-01)
APU GCU
EPAC
GCU 2 GLC2
ALC
BTC2
AC BUS 1
AC BUS 2
3 PH
3 PH
3 PH
AETC
GSTC
AC GND SVC
RLC
3 PH
AC ESS BUS
TRU 1
TRU 2 3 PH
DC GND SVC TRU ESS
STBYC AC EXTERNAL POWER
IDG 1 (SDS 24-21) (MPP 24-21-01)
MIDDLE AVIONICS COMPT
APU GEN
(SSM 24-42-80)
TRU 1C
TRU 2C
RAT GEN (SDS 24-23)
MIDDLE AVIONICS COMPT
STBY AC BUS (MPP 24-23-01)
DCTC
IDG 2 (SDS 24-22) (MPP 24-22-01)
(SDS 24-21) (MPP 24-21-01)
DC BUS 1 LICCGCU 1 (SDS 24-51)
DC BUS 2 EPM
EC 1
(MPP 24-51-01)
GLC 1
BTC 1
RAT GCU
FORWARD AVIONICS COMPT
APU GCU ALC
EPAC
GLC2
BTC2
AC BUS 1 EICC (SDS 24-51) (MPP 24-51-05)
GCU 2 EC 2
AC BUS 2
RF 2 3 PH
EF 3
ETC 1
3 PH
ETC 2
DC ESS BUS 1
AETC
DC ESS BUS 3 EF 2
BC2
GSTC
BC 1
AC GND SVC
3 PH
DC ESS BUS 2
EF 1
RICC (SDS 24-51) (MPP 24-51-03)
RLC
3 PH
AC ESS BUS
TRU 1
TRU 2 3 PH
DC GND SVC STBYC
TRU ESS
TRU 2C
TRU 1C STBY AC BUS
DCTC DC BUS 1 LICC (SDS 24-51) (MPP 24-51-01)
EICC (SDS 24-51) (MPP 24-51-05)
DC BUS 2
EC 1
EC 2
RF 2 EF 3
ETC 1
ETC 2
DC ESS BUS 1
DC ESS BUS 3 EF 2
DC ESS BUS 2 EF 1 BC2
BC 1
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RICC (SDS 24-51) (MPP 24-51-03)
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24-21 Intergrated Drive Generators (IDGs) Introduction Two Integrated Drive Generators normally provide the source of aircraft electrical AC power on ground during taxi, during takeoff and during in-flight operation. Each mechanically-driven IDG is mounted on an engine gearbox. The IDG supplies 30/40 kilovolt-amps at 115/200 volts AC, using a three-phase, brushless type, four output wire system which is stabilized at 400Hz frequency. Stabilized operation frequency is accomplished by the Constant Speed Drive, CSD, which is part of the IDG unit. The purpose of the CSD is to ensure constant rpm by converting variable input speed into a constant output speed by means of a hydro-mechanical mechanism. The constant speed of the generator is necessary to produce the required stabilized 400Hz operation frequency of the AC electrical power supply system. Control and monitoring is provided by the associated Generator Control Unit (GCU), linked to the Secondary Power Distribution Assembly (SPDA) by digital interface (ARINC 429).
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Page 1
Figure 1: IDG Integrated Drive Generator IDG 1
IDG 2
TAXI
TAKE OFF/LANDING IN FLIGHT
CSD variable input speed
GCU constant output speed stabilized 400Hz operation frequency of the AC electrical power supply system
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The Constant Speed Drive The Constant Speed Drive enables the generator to operate at a constant speed of 12,000 rpm, regardless of the variable engine speed and generator load. The input shaft of the CSD drives the generator through a 1:2 ratio gearing and a planetary differential gear. As long as the input drive shaft is at 6000 rpm, the generator operates at 12,000 rpm and produces 400Hz. When the input shaft rotation speed varies, the generator frequency also varies. The IDG is able to achieve this constant Generator drive speed if the input speed is between 4618 rpm,which is the minimum idle speed, and 8130 rpm. This varying speed deviation from 400Hz generates, via the Permanent Magnet Generator (PMG), associated electrical data to the Generator Control Unit (GCU). The GCU supplies an electrical signal to a servo valve inside the CSD. This servo valve then controls - through a control cylinder - the hydraulic log unit, which operates in a clockwise or counter-clockwise direction. Because the hydraulic log unit is connected to another input of the planetary differential gear, the generator rpm increases or decreases less than the input shaft rpm. As a result, the generator frequency stays constant.
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Figure 2: IDG Internal PRESSURIZATION VENT VALVE
GCU
HOLDING TANK
TO SUMP
ECTRONIC NTROLLED DRAULIC RVO VALVE
CYLINDER
ROTATING DEAERATOR
OIL-IN BOSS FROM COOLER
TO SUMP
TO SUMP CHARGE PUMP
CHARGE PRESSURE MONITORING BOSS
CSD
CHARGE PRESSURE SWITCH
HYDRAULIC LOG
CHARGE RELIEF VALVE
TO SUMP
GENERAL COOLING AND LUBE OIL GENERATOR
STATOR
VARIABLE DISPLACEMENT HYDRAULIC UNIT
ACCESSORY DRIVE GEAR
FIXED DISPLACEMENT HYDRAULIC UNIT
DISCONNECT SOLENOID
ROTOR
IMPUT SHAFT CCW ROTATION
PLANETARY DIFFERENTIAL
constant speed of 12,000 rpm
ROTOR STATOR
PMG
GENERATOR SCAVENGE PUMP
SUMP SCAVENGE PUMP
INVERSION SCAVENGE PUMP OIL-OUT BOSS
SCAVENGE FILTER
OVERFLOW DRAIN
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TEMP SENSOR
CASE DRAIN
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PRESSURE FILL
TO COOLER
DPI
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170/190 MAINTENANCE TRAINING MANUAL
The hydraulic log unit
The scavenge relief valve
The hydraulic log unit operates at 240-280 psi pressure. This oil pressure is maintained by a charge pump and charge relief valve. The charge oil is also used to cool and lubricate the generator and the other CSD components. It is not part of the engine oil circuit. The oil flow is between 7 and 12 gal / min. Drain oil from the generator and sump is supplied through a scavenge filter to an external oil cooler by three pumps, which are driven by the accessory drive gear on the planetary differential gear at a constant speed. These pumps are the generator scavenge pump, sump scavenge pump and the inversion scavenge pump. The generator scavenge pump collects lubricating oil from the generator and returns it to the CSD scavenge system to supply the sump scavenge pump. The sump scavenge pump collects the lubricating oil that returns to the bottom of the CSD (sump). These pumps press the oil through the scavenge filter, out of the IDG through the external oil cooler and back into the IDG.
A scavenge relief valve is provided for filter bypass and external circuit when the filter or external circuit remains plugged. In this case the bypass oil flows directly from the scavenge pump to the deaerator inlet. The oil is still circulating within the IDG; therefore the moving mechanism still receives adequate lubrication. However, the oil is not cooled when the IDG is in the bypass mode. The bypassing, uncooled oil will cause the IDG oil temperature to increase, and may cause the amber "FAULT" light in the IDG control panel switch to illuminate. The scavenge relief valve also opens during cold start conditions, because of the high pressure created by high oil viscosity in the external circuit.
Rotating deaerator
The inversion scavenge pump acts as a sump pump during negative-G operation to ensure a continuous flow to the oil cooler and to supply the charge pump.
The entering oil passes through a rotating deaerator. This device extracts air from the oil with centrifugal force. The deaerated oil exits the deaerator discharge and enters the charge pump. The overflow oil leaving the deaerator that does not feed the charge pump fills a holding tank inside the CSD housing.
A Differential Pressure Indicator (DPI)
The holding tank is an all-attitude reservoir designed to assure a continuous supply of oil charge and lubrication oil through all flight attitudes, including negative-G conditions.
A scavenge oil filter is provided between the scavenge pumps and the external oil cooler. This filter cleans the oil before it exits the IDG to prevent contamination of the external oil circuit. A Differential Pressure Indicator (DPI) detects a plugged filter, which means it monitors the oil pressure before and after the scavenge oil filter. In this case the red DPI button will extend as a visual indication.
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Figure 3: A Differential Pressure Indicator (DPI) PRESSURIZATION VENT VALVE
HOLDING TANK
TO SUMP
ELECTRONIC CONTROLLED HYDRAULIC SERVO VALVE
CONTROL CYLINDER
ROTATING DEAERATOR
OIL-IN BOSS FROM COOLER
TO SUMP
TO SUMP CHARGE PUMP
CHARGE PRESSURE MONITORING BOSS
CHARGE PRESSURE SWITCH
HYDRAULIC LOG
CHARGE RELIEF VALVE
TO SUMP
GENERAL COOLING AND LUBE OIL GENERATOR
STATOR
VARIABLE DISPLACEMENT HYDRAULIC UNIT
ACCESSORY DRIVE GEAR
FIXED DISPLACEMENT HYDRAULIC UNIT
DISCONNECT SOLENOID
ROTOR
IMPUT SHAFT CCW ROTATION
PLANETARY DIFFERENTIAL ROTOR STATOR
PMG
GENERATOR SCAVENGE PUMP
SUMP SCAVENGE PUMP
RELIEF VALVE
OVERFLOW DRAIN
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TEMP SENSOR
CASE DRAIN
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INVERSION SCAVENGE PUMP SCAVENGE FILTER
OIL-OUT BOSS
PRESSU FILL
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170/190 MAINTENANCE TRAINING MANUAL
IDG 1 and 2 Selector Knobs AUTO: allows automatic operation of the electrical system. This position closes the IDG contactor, connecting the IDG to the respective AC BUS. OFF: opens the IDG contactor isolating the IDG from the respective AC BUS. DISC: must be held in this position for one second to mechanically disconnect the IDG.
NOTE: an amber LED illuminates indicating to the pilot which IDG must be disconnected.
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Figure 4: IDG Control
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IDG disconnection In case of malfunction, the IDG can be manually disconnected through a switch located on the cockpit control panel (CCP), controlling the electricmechanical disconnect mechanism, which is part of the IDG input shaft. This mechanism consists of a solenoid-operated, spring loaded disconnect plunger, camshaft and reset ring. The manual disconnect should be performed if the associated cockpit - (IDG DISC) amber indicator light and/or the (IDG OIL) CAS message is displayed. These indicate low oil pressure or high oil temperature in the IDG.
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Figure 5: IDG Disconnect
DISCONNECT SOLENOID
ELEC AC SYSTEM
SC
C
IDG 2 CONTROL
IMPUT SHAFT CCW ROTATION AC BUS 2 GPU AVAIL
1 OPEN
AUTO
APU GEN
2 OPEN
IN USE
ELEC DC SYSTEM TRU 1
TRU ESS
TRU 2
DC BUS 1
ESS DC BUS
DC BUS 1
BATT 1 ON OFF
BATT 2 DC BUS TIES
AUTO OFF
The manual disconnect should be performed if the associated cockpit -(IDG DISC) amber indicator light and/or the (IDG OIL) CAS message is displayed.
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170/190 MAINTENANCE TRAINING MANUAL
IDG cautions In case the operational oil pressure in the IDG drops below 140 +/-25psi, the charge pressure switch inside the CSD closes. The GCU interprets this as a low oil pressure condition. In case oil temperature in the IDG sump, sensed by the temperature bulb reaches 335 degrees F (168 degrees C) an IDG over temperature condition is interpreted by the GCU. In this case the GCU will send a corresponding signal over ARINC 429 to the associated SPDA. The SPDA illuminates an AMBER lamp at the IDG DISC switch.
If the IDG oil temperature at the thermal disconnect assembly reaches 366 degrees F, 185 degrees C, the solder pellet melts, and the thermal disconnect pin retracts. The same chain of events occurs during a manual disconnect. However, if the IDG thermally disconnects, pulling on the reset ring will not reset the IDG. The IDG must be returned to the repair shop.
This lamp illuminates if the following SPDA input signals are active: • IDG over temperature signal or low oil pressure signal, and • IDG not already disconnected, and • IDG input speed greater than 4500 RPM, which senses that the engine is running.
Crew actions When the lamp illuminates, the flight crew should take action to hold the IDG selector knob in the “DISC” position for one second. The disconnect mechanism inside the CSD provides a means of separating the transmission and generator shaft from the IDG input shaft.
Thermal disconnect mechanism For safety reasons automatic disconnection is also possible by a thermal disconnect mechanism, part of the IDG input shaft. This mechanism consists of a eutectic solder pellet in combination with a spring loaded retraction pin system. Issue:June06 Revision: 00
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Figure 6: IDG Disconnect
ELEC AC SYSTEM IDG 1 CONTROL DISC
DISC
AC BUS 1 GPU AVAIL
DISCONNECT SOLENOID
IDG 2 CONTROL
IMPUT SHAFT CCW ROTATION
AC BUS 2 AC BUS TIES AUTO
1 OPEN
APU GEN
2 OPEN
IN USE
ELEC DC SYSTEM TRU 1
TRU ESS
TRU 2
DC BUS 1
ESS DC BUS
DC BUS 1
BATT 1 ON OFF
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AUTO OFF
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170/190 MAINTENANCE TRAINING MANUAL
The disconnect solenoid Actuating the disconnect solenoid by turning the cockpit knob results in the solenoid pin retracting. This action releases a plunger, which engages the cam on the transmission shafts. As the shaft rotates, the plunger rides along the cam, moving the shaft axially away from the input shaft. The IDG input shaft part remains engaged with the drive spline of the gearbox. When the IDG DISC knob is turned, a signal is also passed over ARINC 429 to the respective GCU to trip the Generator Control Relay (GCR) and the Generator Line Contactor (GLC).
IDG resetting The disconnect mechanism can be reset by pulling the reset ring on the outside of the IDG. This can only be done when the input shaft is not spinning. The mechanism should only be reset for a disconnect test and not for the reset of an anomaly condition. If the IDG is disconnected for reason, it should be replaced and returned to the repair shop.
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Figure 7: The Disconnect Reset Ring
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170/190 MAINTENANCE TRAINING MANUAL
GCU and PMG For control and regulation, there are inter link circuits between the brushless IDG and the Generator Control Unit. These systems are used to stabilize the IDG output voltage and frequency, and protect the AC generation and supply. A Permanent Magnet Generator, PMG, which is part of the Planetary Differential inside the CSD, rotates with the stabilized IDG speed. The PMG induces three phase 100V / 1200 Hz AC inside the PMG stator, which is sent to the Generator Control Unit. The Power Unit of the GCU rectifies the AC from the PMG into DC (Direct Current) for the Voltage Regulator of the GCU. This regulated DC voltage supplies the ten pole stator windings of the generator`s exciter field. A three phase AC voltage is generated in the exciter rotor by induction. This AC voltage needs to be converted to DC voltage by a rectifier unit which is installed on the generator's common rotor shaft. This DC voltage is supplied to the windings of the four pole main field rotor in which the DC current flow creates a magnet field. This rotating magnetic field induces the output AC voltage in the windings of the generator's main stator. This generator output is connected to the terminal block on the IDG housing. In summary, stabilization of the IDG is accomplished when the generator output AC voltage is stabilized by varying the exciter input DC voltage by the GCU, according to the generator output frequency. The generator output frequency is stabilized by the Constant Speed Drive using the AC voltage and frequency data provided by the Permanent Magnet Generator.
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Page 15
Figure 8: Generator Schematic
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170/190 MAINTENANCE TRAINING MANUAL
NOTES:
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Figure 9: Generator Schematic
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170/190 MAINTENANCE TRAINING MANUAL
24-22 APU AC generation
All these rotor parts are installed on a common shaft, and driven by the APU gearbox.
Introduction
The stator side of the AUX GEN consist of:
The APU system is used primarily when the airplane is on ground for maintenance and flight preparation. The APU generation of AC power can also be used to dispatch the airplane with an altitude restriction or as a backup source of electric power in flight. The APU AC generation system consist of the following main components: • • • • • •
The Generator Unit (AUX GEN), the Generator Control Unit (AUX GCU), the Line Contactor (ALC), the Current Transformer (AUX GEN CT), the APU GEN switch, the crew alerting system (CAS) which is part of the EICAS indication.
• a permanent magnet generator stator, • an exiter stator, • a main stator. All stator coils are of the three phase concept and are installed in the generator housing. Also the three Control Transformers (CT) are installed in these generator housing. The coils of the Control Transformers are connected on the neutral side of the main stator windings. The transformer signals are used by the Auxiliary Generator Control Unit to monitor current for the differential fault detection and protection.
The APU Auxiliary Generator is a four pole, three-phase, brushless type, spray oil cooled and lubricated, rotating rectifier machine. It is rated at 30/40 kVA, 115/200 VAC, 400 Hz. Constant 400 Hz frequency AC power is obtained by rotating the generator with a constant APU engine speed of 12000 RPM. The AUX GEN rotor unit consist of: • • • •
a permanent magnet generator, an exiter field rotor, a diode rectifier assembly, and a main field rotor.
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Page 1
Figure 1: The APU auxiliary generator
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170/190 MAINTENANCE TRAINING MANUAL
The AUX generator The AUX generator is installed on the gearbox of the Auxiliary Power Unit, which is located in the tailcone of the airplane. The generator unit and a seal plate are mounted to the gearbox via threaded studs with washers and nuts. Two alignment pins are installed for ease of installation and for preventing torque motion of the generator housing. The three feeder cables and the neutral cable are attached to the main terminal block of the generator. The terminal blocks of IDG and AUX Gen are the same type. The external connector is used for Permanent Magnet Generator (PMG) stator AC output to the Generator Control Unit (GCU), for Exiter Field DC input from the GCU and for Control Transformer AC signal output to the GCU. For safety purposes, the generator is protected by an input shaft shear section and a thermal disconnect operation. The Auxiliary generator spray lubrication system shares the oil cooling system with APU engine.
The APU generator switch The APU GEN switch is located on the ELEC Control panel, part of the overhead panel. The switch has a latched IN-, and an unlatched OUT-position. A status lamp is part of the swich-button. Latched IN is the usual position, which operates the Auxiliary Generator automatically, and provides power to the aircraft AC distribution system. The status lamp of the switch is OFF in this position. The unlatched OUT position gives the flightcrew the option to manually de-energize the AUX Generator. In this case it opens the Generator Control Relay (GCR)and the Auxiliary Line Contactor (ALC).The switch status lamp is ON in this position.
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Figure 2: The APU generator control
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170/190 MAINTENANCE TRAINING MANUAL
The Auxiliary Generator Control Unit
GCU’s
The Auxiliary Generator Control Unit is a Line Replaceable Module (LRM), it is located in the Right Integrated Control Center (RICC).
The following are maon functions performed by each GCU depending on its location on the aircraft :
The AUX GCU is a microprocessor controlled assembly that gives:
GCU1 (installed in the LICC) - Voltage Regulation and Frequency Control for IDG1 - Protection for IDG1 and its feeders - Control and Protection for AC BUS 1 - No Break Power Transfer (NBPT) for AC System
• • • • •
control, protection, voltage regulation, Generator frequency control and Built-In Test functions.
It will command the Auxiliary Line Contactor (ALC) to open and stop AC power supply to the AC system if a system fault occurs. The AUX Generator Control Unit gives protection to the APU AC-generation system as follows: • • • • • • • • • • • •
overvoltage; undervoltage; overfrequency; underfrequency; overcurrent; phase sequence; differential fault; shorted internal wiring; shorted rotating diode system; inadvertant paralleling trip; open phase; failure of the Central Processing Unit (CPU) of the GCU.
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GCU2 (installed in the RICC) - Voltage Regulation and Frequency Control for IDG2 - Protection for IDG2 and its feeders - Control and Protection for AC Bus 2 - NBPT for AC System AGCU (installed in the RICC) - Voltage regulation for the APU GEN - Protection for the APU GEN and its feeders - Send speed command signal to the APU FADEC for NBPT between the APU GEN and AC EXT PWR - NBPT for AC sytem on the ground - Bus controller for the Inter-LRM communication link between the GCUs and the EPM.
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Chapter 24-22
Page 5
Figure 3: The Auxiliary Generator Control Unit
vervoltage ndervoltage verfrequency nderfrequency
The AUX GCU is a microprocessor-controlled assembly that provides the following generator control, protection, regulation and built-in test functions:
vercurrent hase sequence ifferential fault
RICC
horted internal wiring horted rotating diode system advertent paralleling trip pen phase and failure of the Central Processing Unit (CPU) of the GCU CU
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170/190 MAINTENANCE TRAINING MANUAL
The Auxiliuary Generator Line Contactor (ALC) The Auxiliuary Generator Line Contactor (ALC) is a Line Replaceable Module (LRM) located in the Right Integrated Control Center (RICC). The ALC is a normal open, electrically held contactor that is controlled by the Auxiliary Generator Control Unit. When the AUX GCU has determined that the Aux GEN power quality is good, it sends a signal to close the ALC, then the Auxiliary Generator AC power is supplied to AC BUS 1 and AC BUS 2. The ALC has six usually open and six usually closed auxiliary contacts. The ALC contacts are rated at a minimum continous 150 Amps current per phase. The contactor coil is operated by +28 VDC controlled by the GCU. The pull in current of the coil is less than 2 Amps and a hold current of less than 0.4 Amps.
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Chapter 24-22
Page 7
Figure 4: Auxiliary Generator Control Unit (AUX GCU)
RAT GEN
GLC 1 120A
EICC
Auxiliary generator power 115/200 VAC 400 Hz
GCU
RLC 60A
APU GEN
EPAC 120A
CB3 50A
STANDBY AC BUS TRU ESS 300A
CB4 35A TRU 1 300A
DC GND SVC
TRU EC 400A
TRU 1C 400A EC 1 120A
DC ESS BUS3 EF1 EF2 150A 150A
TRU2 300A
DC BUS 1
Auxiliary Generator Control Unit (AUX GCU) ETC1 120 A
EF3 200A DC ESS BUS 1 BC 1 200A
CB17 15A
EF4 225A
DCTC 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
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RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
AC GND SVC
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
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170/190 MAINTENANCE TRAINING MANUAL
The Auxiliary Generator Line Current Transformer The Auxiliary Generator Line Current Transformer (AUX GEN Line CT) is a Line Replaceable Module (LRM) and is located in the Right Integrated Control Center (RICC). There is one CT for each AC phase. The purpose of each current transformer is to provide differential protection for the phase output leads of the Auxiliary Generator. The Current Transformer (CT) has a transformation ratio of 500:1,used to detect differential fault currents.
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Chapter 24-22
Page 9
Figure 5: APU Generator Block Diagram
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170/190 MAINTENANCE TRAINING MANUAL
APU GEN Electrical The APU GEN is a single-bearing, three-stage, brushless, oil cooled machine. The generator rotor assembly has an exciter rotor, permanent magnet generator (PMG), main generator rotor, and diode rectifier assembly all mounted on a common shaft. The exciter stator, PMG stator and main generator stator are mounted in the APU GEN housing. The generator is spray-oil cooled with oil provided and scavenged by the APU. The APU provides the mechanical power to drive the APU GEN rotor at a nominal speed of 12,000 rpm. The rotation of the PMG induces an alternating current (AC) voltage in the three-phase windings of the PMG stator armature. This AC voltage is supplied to the generator control unit (GCU) where it is conditioned and rectified into DC voltage. The rectified DC voltage is used by the GCU’s voltage regulator to control the current supplied to the windings of the exciter generator field, also called the exciter stator. The stationary magnetic field created by the DC voltage in the exciter stator induces a three-phase AC voltage in the rotating windings of the exciter generator armature (exciter rotor). The rotating diode assembly in the APU GEN rotor assembly then rectifies the AC voltage of the exciter rotor to DC voltage. This DC voltage is applied to the field windings of the main generator rotor. Current flow in the main generator field windings causes a rotating magnetic field, which induces an AC voltage in the main generator stator armature. The power output of the main generator stator is fed through the terminal block on the APU GEN housing, out to the point of regulation. If system conditions are acceptable, the GCU will close the ALC to distribute the APU GEN power to the aircraft loads. Current transformers are mounted on the APU GEN output and monitored by the GCU in combination with the line current transformers at the input side of the ALC for differential protection purposes. To ensure isolation of the APU GEN from the APU gearbox during a high torque generation failure (typically a bearing failure), the drive shaft of the APU gearbox incorporates a tapered section engineered to fail at 533 +/ - 48 inch pounds (this is equivalent to 2100 +/ - 187 inch pounds at the APU GEN shaft).
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Chapter 24-22
Page 11
Figure 6: APU Generator Electrical Schematic
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The EICAS The Engine Indication and Crew Alerting System (EICAS) display screen is located in the center instrument panel. The Crew Alerting System (CAS) provides the pilots with displayed alerts. The CAS display is integrated as a upper right window on the EICAS display. The APU GEN icon is shown in green if the Auxiliary Generator output is greater than 90 VAC, and the APU GEN switch is latched to the IN position. The APU GEN icon is shown in white if the output voltage is less than 70 VAC or the APU GEN switch is unlatched in the OUT position. The APU GEN voltage-number indication shows the output voltage in VAC measured at the Point Of Regulation (POR). The APU GEN frequency-number indication shows the output frequency in Hertz (Hz). The kVA number indication shows the electrical output load in kilo-Volt-Ampère (kVA) always measured at the same point of regulation. AC BUS 1 and AC BUS 2 icons are shown in green if the output voltage is higher than 90VAC, the APU GEN is switched to the IN position, the Aux Line Contactor (ALC) is closed and both Bus Tie Contactors (BTC) are also closed. These AC BUS 1 and 2 icons appears in white, in case output voltage is below 70VAC,the APU GEN switch is in the OUT position,or either Bus Tie Contactor 1 or 2 are open.
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Chapter 24-22
Page 13
Figure 7: Indications
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24-23 AC Emergency Generation Introduction The Emergency AC Generation System provides emergency electrical power for the airplane in the event of a loss of all other sources of AC electrical power. These AC emergency system supplies the AC ESS BUS.
The turbine assembly consist of two turbine blades connected to a mechanical governor, installed inside a hub housing assembly which is connected to the generator shaft.
It consist of the following main components:
The outside diameter of the turbine blades is 24 inches (483 mm). The mechanical governor maintains the rotating speed of the turbine within 7200 and 8800 rpm by the automatic-mechanical regulated variable pitchangle of the the turbine blades. This rotation speed depends upon the the airplane airspeed, altitude and electrical load.
• The Ram Air Turbine, • the the Integrated Control Center and • the associated Generator Control Unit. In case of loss of AC electricity in flight, the Ram Air Turbine AC power will provide the following airplane systems:
The Air Driven Generator assembly is a three-phase,air cooled, brushless AC machine. It produces 15 kVA continous electrical power at 115/200 Volts, and 360 and 440 Hz.
• The AC driven hydraulic pump for the primary flight controls and the landing gear, • the operation of the essential lighting system, • the operation of the essential avionics and communication equipment.
The Ram Air Turbine (RAT) system The Ram Air Turbine (RAT) system, also called the Air Driven Generator (ADG) is installed in a bay in the aircraft nose-right side section. It provides emergency electrical power for the airplane in the event of a loss of all other sources of AC electric power. In the event of an in-flight loss of electrical power, the RAT is deployed automatically into the airstream surrounding the aircraft. The kinetic energy of airflow across the turbine is converted into mechanical power to drive the integral AC-generator. last update: Dec06
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Chapter 24-23
Page 1
Figure 1: The Emergency AC Generation System
RAT Ejection Jack
RAT with Gen (deployed)
RAT with Gen (retracted)
RAT Uplock
RAT GCU
last update: Dec06
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RAT GCU (RGCU) Installation Location The RAT Generator Control Unit (RGCU) is installed in the Forward Fuselage in a temperature controlled and pressurized region on the right side in the Forward E-Bay of the aircraft. The RAT PMG powers the RAT GCU. No aircraft power is required. The RAT GCU has the following functions: - Provides RAT Generator Voltage Regulation - controls RAT Line Contactor (RLC) allowing RAT GEN power to supply AC ESS BUS - Provides Overvoltage protection (sensed at POR) - Trips RLC - Automatic reset if voltage subsequently OK - Provides Under Frequency protection (PMG sensing) - Trips RLC - Automatic reset if frequency subsequently OK - Provides BITE function - Tests Overvoltage circuit$ - cycles RLC to test drive circuit - BIT Pass indicated by LED illumination - BIT Fail indicated by LED not illuminates - RLC will not close - Provides HIRF / Lightning protective wire shielding
last update: Dec06
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Chapter 24-23
Page 3
Figure 2: RAT GCU Location
GCU Ac c e s s Do o r
last update: Dec06
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170/190 MAINTENANCE TRAINING MANUAL
RAT deployment The Ram Air Turbine (RAT) deployment occurs automatically in case of an airborne loss of AC power from both Integrated Drive Generators (IDG1 and IDG2). The Ram Air Turbine can also be deployed manually by a flight crew member using a deployment lever located in a console between the flight crew seats. The distribution of emergency AC power is controlled and monitored through contactors, circuit breakers, relays and protections which is provided by the RAT-Generator Control Unit (RATGCU) and the Emergency Integrated Control Center (EICC). The main stator in the generator of the RAT has an electrical heater installed to prevent moisture from freezing in the air gap between the stators and rotor during cold temperatures and icing conditions. This feature helps assure smooth startup at deployment. The RAT will remain deployed and operational during the entire flight and the landing phase. After a RAT system inspection procedure, maintenance technicians can restow the RAT with the restow pump.
last update: Dec06
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Chapter 24-23
Page 5
Figure 3: RAT deployment
last update: Dec06
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RAT system Components The RAT System is comprised of the following components : - RAT with Generator - Mechanically governed turbine assembly, which direct drives a 3-phase brushless generator to produce 115 volts/400 Hz power at the Point of Regulation (POR) - RAT Actuator - Spring-loaded, hydraulic actuator, which provides the initial forces required to deploy the RAT - Also provides hydraulic damping to limit impact loads at the end of deployment - RAT Restow Pump - Utilized to manually retract the RAT to the stowed position by providing hydraulic pressure to the Ejection Jack - RAT Uplock - Holds RAT in stowed position until release is commanded - RAT GCU - Provides excitation control for the RAT electrical output - Monitors power quality of RAT electrical output for system protection purposes - Coordinates closure of RLC, allowing RAT GEN to power AC ESS BUS
last update: Dec06
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Chapter 24-23
Page 7
Figure 4: RAT System
The three-phase AC generator produces 5 kVA 400 Hz stabilization is regulated by the automatic-mechanical variable pitch angle of the turbine blades.
15/200 V 00 Hz stabilized.
RAT
last update: Dec06
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RAT System functional Interface If both IDG1 and IDG2 fail with the APU GEN not available, each GCU commands its respective GLC/ALC tripped open. Under these conditions, SPDA logic recognizes, via system configuration contactor command status over ARINC 429, that no generator is operating. If in air mode with appropriate air speed, the SPDA automatically commands the RAT to be deployed. As a backup, there is a mechanical lever available in the cockpit to manually deploy the RAT is necessary. During this emergency mode operation, AC ESS BUS load logic is implemented to ensure RAT loading is properly coordinated. After the RAT GCU senses that the RAT has enough governing speed to power loads, the RAT GCU closes the RLC, allowing the RAT to power the AC Motor Pump 3A. DC ESS BUSes remain powered via Batteries during this transition. The timing of the DC ESS BUS loads and the AC ESS BUS Motor Pump 3A does not overload the RAT. The 1-second delay coordinates control of the TRUEC contactor to delay RAT power being applied to the DC ESS BUSes. After the RLC closes, DC ESS BUS 1 power from BATT 1 is routed to 1-second delay relay. After the time delay has been met, the 1-second delay relay closes. This allows the TRUEC to close. During landing, the MAU removes a ground for the TRUEC enable relay as air speed is sensed below 160 knots. This causes the TRUEC to open and allow the RAT to supply the AC ESS BUS (AC Motor Pump 3A) independent of the DC ESS BUSes, which are then powered by the batteries. With the BATT 1 switch in the ON position and the BATT 2 switch in the AUTO position, BC1 and BC2 are closed. With the TRU ESS and DC BUS TIES switches latched IN position, TRUEC, ETC1, ETC2 are commanded closed to ensure that battery charging takes place during RAT deployment.
last update: Dec06
The TRU ESS converts the three phase AC input power from the RAT into a + 28 VDC output to supply the DC ESS Bus 1, DC ESS Bus 2, and DC ESS Bus 3. BATT 1 and BATT 2 receive a charging current through the associated BC1 and BC2 contactors if the BATT1 switch is in the ON position and the BATT 2 switch is in the AUTO position.
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Chapter 24-23
Page 9
Figure 5: RAT System
RAT GEN
IDG 1
AC EXT PWR
APU GEN
IDG 2
EICC F1 50A 0
TC 2 20A 0
CC
225A 5 HOT O BA BATT BUS 2 ABC 400A 0
AF2 AS A C 300A 0 400A 0
EPDC 400A 0
TO APU T STA T RT R DC EXT PWR
BA B ATT 2
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170/190 MAINTENANCE TRAINING MANUAL
NOTES:
last update: Dec06
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Chapter 24-23
Page 11
Figure 6: RAT System Schematic
FORWARD AVIONICS COMPARTMENT
MIDDLE AVIONICS COMPT
MIDDLE AVIONICS COMPT
AC BUS 1
AC BUS 2
A
A
B
B
C
C
LICC (SSM 24-51-80)
RICC (SSM 24-51-80)
FORWARD FUSELAGE
AETC
RAM AIR TURBINE GENERATOR
G
FORWARD FUSELAGE (SDS 24-23) (MPP 24-23-01)
PHASE C PHASE B PHASE A
RAT HEATER
POR
PMG
EXCITER FIELD
B HEATER PWR
(SDS 24-23) (MPP 24-23-05)
(SDS 24-51) (MPP 24-51-11)
C
RAT DEPLOY SOLENOID
H
RLC
J
(SDS 24-23) (MPP 24-23-13)
RLC CMD MONITOR RAT GCU TEST HEATER PWR FWD AV COMPT
FWD FUSELAGE AC ESS BUS A SPDA 1 (SSM 24-61-80)
last update: Dec06
RAT GENERATOR CONTROL UNIT (SDS 24-23) (MPP 24-23-09)
B
E
EICC (SSM 24-51-80)
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C
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170/190 MAINTENANCE TRAINING MANUAL
NOTES:
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Page 13
Figure 7: RAT System
last update: Dec06
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170/190 MAINTENANCE TRAINING MANUAL
24-24 Static inverting Introduction The Static Inverter is located in the temperature and pressure controlled Forward E-Bay. Access to the Static Inverter can be accomplished through the Forward E-Bay floor access hatch which is located in front of the nose gear. The static inverter is a Line replaceable Unit (LRU). The purpose of the unit is to convert 28 volts DC into AC,in order to provide the Standby AC Bus with single-phase, 115 volts RMS (Root Mean Square), 400 Hz AC stabilized power output. The capability to convert DC into AC is 250 VA (Volt-Amps) of electrical power.
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Chapter 24-24
Page 1
Figure 1: The Static Inverter
FWD E-bay Static inverter
STANDBY AC BUS
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AC
AC INVERTER 250VA DC
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28 VDC
Chapter 24-24
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Operation During operation input of the static inverter ist powered by DC Essential Bus 1 (ESS BUS 1), this happens in case power is not available to the AC Essential Bus (AC ESS BUS). Then the Standby Contactor (STBYC) inside the EICC will be de-energized, connecting and suppling the AC output of the inverter to the Standby AC Bus. The operation of the inverter is controlled by Secondary Power Distribution Assembly 1 (SPDA1). These unit controls the ON or OFF logic of the static inverter based upon the inverter internal fault monitor and the number of AC sources available during in flight operation. The SPDA1 will set the inverter to the OFF status during normal operation of the Electrical-Power Generation-and-Distribution System (EPGDS). The SPDA1 will set the inverter to the ON status in case only one main AC source is available (IDG 1 or IDG 2 or the APU generator). If the internal fault monitor of the inverter detects a problem, a discrete status signal will inform the SPDA 1 to cause a shutdown of the inverter.
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Chapter 24-24
Page 3
Figure 2: Operation
RAT GEN
GLC 1 120A
EICC RLC 60A
EPAC 120A
CB3 50A
CB26 25A
TRU ESS 300A TRU EC 400A
TRU2 300A
TRU 1C 400A DC BUS 1 EC 1 120A
DC ESS BUS3 EF1 EF2 150A 150A
ETC1 120 A
EF3 200A DC ESS BUS 1 BC 1 200A
CB17 15A
EF4 225A
DCTC 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
Static inverter
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AC INVERTER 250VA DC
FOR TRAINING ONLY Reproduction Prohibited
RICC
AC BUS 2 CB1 CB29 50A 35A
CB4 35A TRU 1 300A
DC GND SVC
GLC 2 120A
BTC 2 120A
GSTC 60A
AC GND SVC
IDG 2
ALC 120A
LICC
AC BUS 1
STANDBY AC BUS
SPDA 1
APU GEN
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
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24-30 DC Generation System Introduction DC power will be generated by: • three Transformer Rectifier Units, • two accumulator batteries, • or an external DC power connection. For operation of DC-powered airplane equipment, the conversion from AC power to DC power is provided by three Transformer Rectifier Units ( TRU`s). Batteries 1 and 2 are used to backup all DC buses to ensure a continous supply for DC loads. Battery 2 also provides energy necessary for an APU start. DC external power is provided via a Ground Power Unit (GPU) and is only used to perform an APU start.
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Page 1
Figure 1: DC power sources
DC power will be generated by:
3 TRUs
Two accumulator batteries
Ground Connector
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24-31 DC Generation Introduction The DC system of the airplane operates by 28 Volts DC stabilized. There are three Transformer Rectifier Units (TRUs) as main DC power sources. The TRUs are located inside the associated Intergrated Control Centers (ICCs). There are also two Nickel-Cadmium batteries and one External DC power plug.
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Page 1
Figure 1: DC Power
DC power will be generated by:
3 TRUs
Two accumulator batteries
Ground Connector
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DC power DC power is generated by conversion of AC to DC by the TRUs to supply the associated DC bus. The input of each TRU is 115VAC, 400Hz frequency stabilized, and the output is 28VDC regulated up to 300 Amps of current. Both Nickel Cadmium batteries have 19 accumulator cells. Battery 1 is located in Forward E-Bay, Battery 2 is located in the Aft E-Bay.
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Page 3
Figure 2: Transformer Rectifier Unit
input
TRU
output
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170/190 MAINTENANCE TRAINING MANUAL
Normal operation In normal operation any available power source will provide charging current for batteries 1 and 2. Both batteries are in operation to backup all DC buses ensuring a break free environment for DC loads. Battery 2 also provides electrical power for an APU start through the APU start bus, when the electrical system is automatically isolated from battery 2. DC external power is routed through the airplane DC power receptacle to the External DC Power Contactor (EPDC) and is used for powering the APU Start Bus upon an APU start attempt. External DC power is controlled by dedicated system relay logic.
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Page 5
Figure 3: APU start
OT SWITCH2 OT SWITCH2 RTN
TRU VOLT SENSE1 TRU VOLT SENSE2
+ INTERPHASE TRANSFORMER
RECTIFIER BRIDGE 1
-
AC PHASE A AC PHASE B AC PHASE C
EMI INPUT FILTER
OUTPUT FILTER / BLEEDER
POWER TRANSFORMER
+ RECTIFIER BRIDGE 2
-
PWR 28V
-
+ SHUNT
PWR GND TRU VOLT SENSE1,2 RTN TRU SHUNT1(+) TRU SHUNT2(+) TRU SHUNT1(-) TRU SHUNT2(-) OT SWITCH1 OT SWITCH1 RTN
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24-36 Main Batteries Introduction Two Nickel-Cadmium (NiCd) Accumulator-Batteries are provided for powering essential loads if Transformer-Rectifier-Unit (TRU) power is not available. Battery 1 and battery 2 are be used to backup all DC buses to ensure a break free environment for DC loads.
Battery case container Each battery consists of a steel case containing 19 semi-open NiCd-cells. The cells are connected to each other in series by copper bus bars. The nominal battery voltage is 22,8 VDC, with a capacity of 27 Ah (Amp-hours, at a 1 hour discharge rate). Adequate battery ventilation is provided through a tube interconnecting the battery-venting nozzle to the fuselage surface. A caliber orifice close to the fuselage and a check valve on the battery vent inlet assure that gases emitted by the batteries, under normal or abnormal conditions, will not accumulate hazardous quantities.
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Page 3
Figure 2: EICC Battery 1 Contactor (BC1)
TRU EC
EF1
EF2
EF4
ETC1
BC1
K5
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K8
K6
K7
EF3
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Battery location Battery 1 is located in the forward E-bay directly below the SPDA 1 unit. Battery 1 provides stored energy to selected equipment during normal operations and during flight in the absence of all other airplane electrical power. Battery 2 is located in the aft E-bay directly below AICC. Battery 2 also provides the power used for an APU start through the APU start bus during ground or flight operations, when the electrical system is automatically isolated from the battery 2.
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Page 5
Figure 3: Battery 2
AFT E-Bay
AICC Battery 2
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Temperature sensors Two temperature sensors, fitted in each battery, are continiously monitored by the Modular Avionics Units (MAU`s). Both sensors are used to provide continuous indication of the battery temperature on the Multfunction Display (MFD) and to display an EICAS warning message "BATT OVERTEMP" if battery limits are exeeded. The higher of the two temperature values sensed is the one utilized for indication and alarm. Whenever a battery overtemperature condition is sensed, the battery should be isolated by the pilot from the charging source. The chemical nickel cadmium cells are protected by the battery case. Battery overtemperature will never generate sufficient heat to damage the battery`s surroundings, including those caused by a short circuit at it terminals or at any of it cells. Any mechanical deformation of the battery will be contained within its steel case.
Maintenance requirements for the battery include a regular check every 600 flight hours and a general overhaul every 12 month.
Battery contactor BC 1 (mounted on the EICC) connects the avionics battery (BATT1) to DC ESS BUS 1 when the BATT 1 switch is in the ON position. This enables BATT 1 to provide DC power to the following when there is no other power source available : - DC ESS BUS 1 - DC ESS BUS 3 when ETC 1 is closed - DC ESS BUS 2 when ETC 1 and ETC 2 are closed - Static Inverter
The battery must be removed from the airplane for both the regular check and the general overhaul. The regular check consists of the following: • • • • • • • •
exterior cleaning, voltage check, nut tightness check, discharge and shorting of all cells, insulation check, recharge, adjustment of electrolyte level, capacity check, cleaning of battery vents.
The general overhaul consists of the following: • • • • • • • • • • •
voltage check, nut tightness check, discharge disassembly and shorting of all cells, thorough cleaning and inspection of all components, check of the thermal sensors, replacement of fault components, assembly of all battery components, insulation check, recharge and adjust the electrolyte level of all cells, capacity check of battery, cleaning of battery vents.
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APU Start Battery (BATT 2) Functional Interface Battery contactor BC 2 (mounted in the RICC) connects the APU start battery (BATT 2) to DC ESS BUS 2 when the BATT 2 switch is in the AUTO position. This enables BATT 2 to provide DC power to the following when there is no other power source available : - DC ESS BUS 2 - DC ESS BUS 3 when ETC 2 is closed - DC ESS BUS 1 when ETC 1 and ETC 2 are closed - Static Inverter
24-40 External Power Introduction The airplane gets its external AC power from the Ground Power Unit (GPU). The three-phase, 115 Volt, 400 Hz external AC power is used for ground maintenance and flight preparation. External DC can be used for APU start. External 3-phase, 115 Volts AC, 400 Hz may be connected to the aircraft via a receptacle on the LH side of the fuselage. The external AC is used to power the AC ground service bus and the DC ground service bus without powering any other busses on the airplane. The operation of the system is controlled by the External Power Module (EPM).
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Page 1
Figure 1: External Power
IN USE
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GPU Power Panel The plugged-in six-pin external power plug activates the External Power Module in order to monitor the quality of the supplied AC. The E-F interlock monitors the proper plug set. The AVAIL lamp on the AC GPU Power Panel will illuminate if power quality is acceptable and system pin E-F interlock is achieved. At the same time, the GPU AVAIL lamp on the Cockpit Control Panel will illuminate. The Ground Service switch on the AC external power panel allows activation of the GSTC from outside the airplane. When the AVAIL indication inside the switch changes into IN USE, the ground service buses will be powered. The Ground Power Unit (GPU) switch in the cockpit provides the ability to activate the system from inside the airplane. When the AVAIL indication inside the switch changes into IN USE, the main Aircraft AC buses will be powered, plus the ground service buses, via AC bus 1.
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Figure 2: GPU Power Panel
AC GPU MIC AVAIL
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EPM Installation One EPM is utilized in the EMB 170/190 EPGDS design. It is installed in the LICC to provide control and protection for the AC EXT PWR channel. Access to the EPM is accomplished through the Mid E-Bay floor access hatch, which is located on the aircraft left side behind the left wing. The EPM is mounted in a pressurized location with no forced air cooling. All cooling of the EPM is through natural convection. The following are the main functions performed by the EPM : EPM (installed in the LICC) - Control and Protection for AC EXT PWR channel - Control of EPAC and GSTC - Control for Pins E/F Interlock
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Figure 3: External Power Module
EPM
LICC
Floor Access Hatch
Mid E-Bay
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EPM (Block Diagram) The EPM primary function is to provide AC external power protection. There is one EPM LRM integrated in the EPGDS and this component is physically located within the Left ICC. This EPM consists of a two Printed Wiring Board (PWB) assembly mounted in a board stack Line Replaceable Module (LRM). The A1 PWB performs the processing functions for the EPM and the A2 PWB is the Input/Output interface with the A1 processing PWB as well as the power supply and power drive for system functions. The A1 PWB contains the digital hardware and related peripherals of the EPM, including the microprocessor which is an 80C186 supplied with a 40MHz clock (instructions run at half this speed, 20 MHz). The memory consists of a 512k x 16 Flash for program (flight code), 32k x 8 Non-Volatile Memory (NVM) for Bite Storage and retrieval and a 128k x 16 RAM for data transfer. A Watchdog Timer (WDT) is also resident to monitor for proper microprocessor operation. If an errant operation is detected, a fail-safe is initiated. Support hardware includes the Application Specific Integrated Circuit (ASIC), the Digital Signal Processor (DSP) and the Analog to Digital Converter (A/ D). The ASIC device assists the microprocessor by performing repetitive tasks such as serial communications (1553 and RS485) control, frequency and phase angle measurements, A/D control, phase sequence protection, chip select and address latching. The DSP provides the ASIC with processed External Power voltage and current data. The A/D converter converts raw analog data to a digital equivalent. Input/ Output buffers provide and receive discrete information to and from the A2 hardware sensing/driver circuits.
nected (DC BUS 1 unpowered), the EPM routes +28 VDC through a driver to the AGCU. The AGCU has BTC control with the respective GCU1 and GCU2 unpowered. Beyond the power supply function, the A2 PWB performs many input/output functions to receive various discrete signals (switches, auxiliary contacts etc.) from the system and provides filtering prior to passing the information to the A1 microprocessor board for processing. An interface for discrete outputs (lamp driver commands etc.) to the external environment and an interface for contactor driver Solid State Relays (SSRs) is provided on the A2 PWB to provide contactor control to energize/de-energize system contactors including the EPAC and GSTC. Analog signals (External Power voltages and currents) are also received through the A2 PWB and filtered prior being processed on the A1 PWB. Additionally, this A2 PWB contains the serial communications bus drivers for the 1553 inter-LRU bus and RS485 test link. All A2 PWB signals are isolated and filtrered in order to provide HIRF and lightning protection. Another function provided through the A2 PWB relates to the Pin E/F interlock circuitry. During operation with an AC external power source connected to the aircraft, the EPM routes a nominal 28Vdc from pin F through an external AC Ground power source and back to pin E to verify appropriate interlock of the external power receptacle. A 28 Vdc nominal voltage is provided on pin F when enabled by the microprocessor. An indication that 28Vdc is present at Pin E is monitored by the microprocessor (A1) to determine if the External Ground Power Plug is properly mated to the aircraft (proper interlock).
The A2 PWB interfaces directly with the 3-phase AC External Power and +28VDC power (from DC BUS 1) to power the EPM. The A2 PWB uses an Internal Power Supply (IPS) to regulate the input power to needed +15 volt, - 15 volt, + 5 volt and filtered + 28 volt power for various A2 PWB and A1 PWB circuits. The switching and filtered power supplies are monitored and if their performance is out of specification, a control unit fail-safe shall be initiated. To allow closure of the BTCs when only external AC Power is con Issue: June06 Revision: 00
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Figure 4: EPM Schematic
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24-41 DC External power Introduction DC EXTERNAL POWER switch is located on the DC External Power Panel. It is a push-button latching switch. This switch provides control of the External-Power-DC-Contactor (EPDC). This is controlled by a power relay logic of the External Power Module (EPM) inside the Auxiliary Integrated Control Center (AICC). If acceptable external DC power quality exists with the DC EXTERNAL POWER-switch in the OFF position, the DC EXTERNAL POWER AVAILABLE lamp, located on the DC External Power Panel, will illuminate via aircraft wiring, and the EXTERNAL POWER IN USE lamp will be OFF. When the External DC Power Source is plugged in, power quality is acceptable and the DC EXTERNAL POWER switch is closed, the External Power DC Contactor (EPDC) will close. With the EPDC closed, the DC EXTERNAL POWER IN USE lamp, located on the DC External Power Panel, will illuminate via aircraft wiring, and the AVAIL lamp will be OFF. An additional indication of DC GPU IN USE will be provided as a Crew Alert System Advisory message. With the DC EXTERNAL POWER switch in the OFF position, the EPDC will open.
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Page 1
Figure 1: DC External Power
EPM
(AC GPU) external AC power
The three-phase, 115 Volt, 400Hz external AC power is used for ground maintenance and flight preparation
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The operation of the system is controlled by the External Power Module (EPM)
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CAS Caution Message GPU CONNECTED SPDA2 generates CAS Caution Message GPU CONNECTED when the parking brake has been released while the AC ground power unit (AC GPU) is still connected to the AC EXT PWR receptacle or the DC groung power unit (DC GPU) is still connected to the DC EXT PWR panel. CAS (MAU) inhibits the message during takeoff, landing, and in flight.
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Figure 2: CAS Caution Message
GPU CONNECTED
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24-42 AC-External Power Cockpit Electrical Panel. The cockpit control panel's Ground Power Unit (GPU ) switch provides control of the external AC power. Once the GPU is connected and power quality requirements are satisfied, a GPU AVAIL indication will illuminate on the GPU switch in the cockpit.. Pressing the GPU switch in the cockpit will allow operation of the EPAC, connecting the external AC power to the main AC buses, and a GPU IN USE indication will illuminate.
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Figure 1: The cockpit control panel
AVAIL
AVAIL
IN USE
IN USE
(AC GPU) external AC power
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AC External Power Panel A momentary action type Ground Service switch is located on the AC External Power Panel. It is marked as AC GPU, and includes the AVAIL and IN USE indicator lights. This Ground Service switch provides control of the Ground Service Transfer Contactor (GSTC) through the External Power Module ( EPM ) inside the Left Integrated Control Center ( LICC ). The GPU AVAIL lamps on the Cockpit Control Panel and on the External Power Panel will illuminate if power quality is acceptable and the interlock system (pin E/F on the Ext.-Power-Plug) is in order. When the External AC Power Unit (GPU) is plugged in, power quality is acceptable, AC BUS 1 is unpowered, and the momentary GroundService switch is toggled, the Ground Service Transfer Contactor (GSTC) will close. The GPU IN USE lamp on the External Power Panel will illuminate with the Ground Service Transfer Contactor (GSTC) closed. This closure of the GSTC results in only the AC Ground Service Bus and the DC Ground Service Bus being powered. If the AC BUS 1 is powered, then the Ground Service Transfer Contactor will open. If the GSTC is energized and the Ground-Service-switch is again toggled, the GSTC will move to the rest position.
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Figure 2: AC External Power Panel
AC GPU MIC AVAIL
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24-51 AC Power Generation
The AC power distribution
Introduction
The AC power distribution supplies AC power to various airplane loads through several AC buses. The AC buses are physically and electrically segregated to be sure that the loss of one AC bus not affect the other.
AC power distribution is performed by subassemblies located in the: • • • •
Left Integrated Control Center (LICC), Right Integrated Control Center (RICC), Emergency Integrated Control Center (EICC) and by Secondary Power Distribution Assemblies (SPDA 1 and SPDA 2).
two main AC-channels, one auxiliary channel, one external power channel and a Ram Air Turbine channel.
Each main AC channel receives power from associated Integrated Drive Generators (IDG 1 and IDG 2) through a Generator Line Contactor (GLC 1 and GLC 2), and distributes AC power via AC bus 1 and AC bus 2. The auxiliary channel receives power from the Auxiliary Power Unit (APU) through an Auxiliary Line Contactor (ALC) and distributes AC power through a Bus Tie Contactor (BTC1 / BTC2) and via AC Bus 1 / AC bus 2. External AC ground power can also be supplied to the airplane through the Alternating Current Contactor (EPAC). The Ram Air Turbine channel receives power from the RAT generator through the Ram Air Turbine Line Contactor (RLC), and distributes AC power via AC essential bus. Issue: June06 Revision: 00
The main buses supply AC power to various switched airplane loads such as: • • • • • •
AC power for the airplane is received from: • • • •
The two main AC buses are the left main AC bus 1 which is located in the Left Integrated Control Center (LICC), and the right main AC bus 2, located in the Right Integrated Control Center (RICC).
the galley feeds, galley heaters, AC fuel pumps, windshield heater control units, airplane lights, AC fans and lavatory heaters.
They also supply AC power to other loads through the Secondary Power Distribution Assembly (SPDA 1 or SPDA 2). The AC Ground Service bus (AC GND SVC) is supplied out of AC bus 1 or external AC source through the Ground Service Tie Contactor (GSTC). The AC Essential bus, located in the Essential Integrated Control Center (EICC), is supplied through Alternating Current Essential Tie Contactor (AETC). The Standby AC Bus is supplied out of the AC Essential bus through the Standby Contactor (STBYC). AC bus 1 or the AC Ground Service bus supplies Transformer Rectifier Unit 1 (TRU 1); AC Bus 2 supplies TRU 2, and the AC Essential Bus supplies Essential TRU (TRU ESS).
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Figure 1: AC power distribution
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
TRU2 300A
TRU 1C 400A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
DC BUS 1
ETC1 120 A
CB17 15A
DCTC 120A
EF3 200A DC ESS BUS 1 BC 1 200A
AICC
EF4 225A
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
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RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
EC 1 120A DC ESS BUS3 EF1 EF2 150A 150A
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
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ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
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24-52 AC Circuit Breakers Introduction The circuit breakers are thermal type devices with temperature compensation. When subjected to an overload current, the breakers will trip open after a predetermined time. The circuit breakers are the free tripping, push/pull, on/off, manual actuation type. The CB aux contacts are normally open, polarized with a blocking diode. The AC circuit breakers (CB) provide protection of supply lines and cabeling for all loads whose power is sourced from the AC supply system of the airplane. The AC and DC circuit breakers are located in the LICC, RICC and in the EICC. On the AICC, only DC CBs are installed. Some fuses are also installed on the front panel of the LICC, RICC and EICC. Additional circuit breakers are located in the flight deck as left and right circuit breaker side panels. Caution! All power sources to the ICCs need to be disconnected prior to performing maintenance on the ICC, such as opening the panels or removing any Line Replaceable Modules (LRMs).
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Figure 1: AC circuit breakers
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
LICC
BTC 1 120A
AETC 60A
CB3 50A
CB26 25A
AC GND SVC
STANDBY AC BUS TRU ESS
IDG 2
ALC 120A
GLC 2 120A
RICC
BTC 2 120A
AC BUS 1
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
GSTC 60A CB4 35A TRU 1
AC BUS 2 CB29 CB1 35A 50A
TRU2
The main buses supply AC power to various switched airplane loads such as
galley feeds lley heaters fuel pumps ndshield heater control units plane lights fans atory heaters
They also supply AC power to other loads through the Secondary Power Distribution Assembly (SPDA 1 or SPDA 2). DC
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The LICC The Left Integrated Control Center provides AC and DC power distribution and protection for the main airplane busses associated with the left side of the airplane. This list shows all AC circuit breakers and their ratings, installed on the front panel of the LICC. The upper AC BUS 1 CB panel can be released by unscrewing the four captive screws and pulling the panel away from the LICC. The wires are long enough to enable replacement of any of the three CBs or two fuses. The lower front access panel of the LICC can be opened after releasing the eleven captive screws, allowing easy access to the circuit breakers and other components inside. Don`t forget to disconnect electrical power from the LICC prior to performing maintenance!
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Figure 2: The LICC
RAT GEN
IDG 1 GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
CB17 15A
TRU 1C 400A
TBD TBD
GLC 2 120A
AC BUS 1
TBD
CB12
FWD (Forward) GALLEY HEATER
AC BUS 1
TBD
CB13
AC FUEL PUMP 1
AC BUS 1
TBD
CB14
AC FEED
AC BUS 1
TBD
CB15
LEFT FAN
AC BUS 1
TBD
CB16
SPDA (Secondary Power Distribution Assembly)2 AC 1 FEED
AC BUS 1
TBD
CB17
WHCU (Windshield Heating Control Unit) 2 MODULE FEED
AC BUS 1
TBD
CB18
GALLEY 2 FEED 1
AC BUS 1
TBD
CB19
GALLEY 3 FEED 4
AC BUS 1
TBD
CB04 CB05
TRU (Transformer Rectifier Unit)1 PWR (Power) SPARE
AC GROUND SERVICE BUS AC GROUND SERVICE BUS
TBD TBD
ISOLATED RICC
CB20
FAN 1 FWD BAY
AC GROUND SERVICE BUS
TBD
CB21
FAN 1 MID BAY
AC GROUND SERVICE BUS
TBD
AF2 ASC 300A 400A
CB22
CEILING LIGHTS
AC GROUND SERVICE BUS
TBD
CB23
SIDEWALL LIGHTS
AC GROUND SERVICE BUS
TBD
CB24
SPDA1 FEED
AC GROUND SERVICE BUS
TBD
CB25
WATER/WASTE HEATER
AC GROUND SERVICE BUS
TBD
CB26
VACUUM MOTOR GEN (Generator) FEED
AC GROUND SERVICE BUS
TBD
CB27
AC OUTLET PWR
AC GROUND SERVICE BUS
TBD
CB28
SPARE
AC GROUND SERVICE BUS
TBD
CB29
SPARE
AC GROUND SERVICE BUS
TBD
RICC
EF3 200A DC ESS BUS 1 BC 1 200A EF4 225A
DCTC 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
CB RATING
AC BUS 1 AC BUS 1
TRU2 300A
DC BUS 1
ETC1 120 A
SOURCE
AC BUS 2 CB29 CB1 35A 50A
GSTC 60A
EC 1 120A DC ESS BUS3 EF1 EF2 150A 150A
NAME HYD (Hydraulic) MOTOR PUMP 2B CBP AC BUS 1 FEED AETC (Alternating-Current Essential Tie-Contactor) POWER FEED
BTC 2 120A
AC BUS 1 CB3 50A
CB26 25A
CIRCUIT BREAKER CB01 CB02 CB03
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
BATT 1
BATT 2
EPDC 400A
TO APU START DC EXT PWR
SLAT 1
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The RICC The Right Integrated Control Center provides AC and DC power distribution and protection for the main airplane busses associated with the right side of the airplane. This list shows all AC circuit breakers and their ratings, installed on the front panel of the RICC. Removal and replacement is done by loosening the attached hardware and the interface wire harness. These devices are not to be reset or replaced in flight.
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Figure 3: The RICC RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
CB17 15A
TRU 1C 400A DCTC 120A
EF3 200A DC ESS BUS 1 BC 1 200A EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A
BATT 1
NAME
SOURCE
CB RATING
CB02
HYD MOTOR PUMP 1B
AC BUS 2
CB15 CB16
FAN 2 MMID BAY AC BUS 2 FLAP 2 AC FEED AC BUS 2
CB17
SPDA 2 AC 2 FEED
AC BUS 2
35 A AC 3 PHASE 5 A AC 3 PHASE 15 A AC 3 PHASE 25 A AC 3 PHASE
CB18 CB19
FAN 2 FWD BAY RIGHT FAN
AC BUS 2 AC BUS 2
5 A AC 3 PHASE
CB20
GALLEY 3 FEED 1/5 WHCU 1 MODULE FEED
AC BUS 2
25 A AC 3 PHASE
AC BUS 2
25 A AC 3 PHASE
CB22
SPARE RLY (RELAY) FEED
AC BUS 2
25 A AC 3 PHASE
CB23
GALLEY 2 FEED 3 AFT GALLEY HEATER
AC BUS 2
25 A AC 3 PHASE 5 A AC 3 PHASE
RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
AC CIRCUIT BREAKERS
TRU2 300A
DC BUS 1
ETC1 120 A
RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
EC 1 120A DC ESS BUS3 EF1 EF2 150A 150A
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 2
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
CB21
CB24
AC BUS 2
CB25
VIBRATION SYSTEM
AC BUS 2
5 A AC 3-PHASE
CB26
DOOR SILL HEATER
AC BUS 2
7.5 A AC 3-PHASE
CB27
CBP AC B S 2 FEED HYD MOTOR PUMP 3B TR 2 PWR
AC BUS 2
35 A AC 3-PHASE 35 A AC 3-PHASE 35 A AC 3-PHASE
LOAD BANK FEED
AC BUS 2
CB28 CB29
AC BUS 2 AC BUS 2
Right Integrated Control Center RICC CB35
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The EICC The Essential Integrated Control Center provides AC and DC power distribution and protection for the airplane essential, standby, and hot battery buses. This list shows all AC circuit breakers and their ratings, installed on the front panel of the EICC. The first list shows seven CBs and 2 spares, associated with the AC Essential Bus (AC ESS BUS). The second list shows two CBs associated with the STANDBY AC BUS. This shows the AC Essential Bus CBs, and the Standby AC bus CBs. There is one front panel on the EICC. Releasing eleven captive screws provides easy access to the rear of the CB installation. Reminder: Take precautions to ensure proper power and Electro Static Discharge (ESD) safety considerations while performing maintenance on the open ICCs. Electrical power must be disconnected an ESD wrist strap must be worn since there are static sensitive modules installed in the ICCs.
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Figure 4: The EICC RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB3 50A AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
TRU2 300A
TRU 1C 400A DC BUS 1
ETC1 120 A
CB17 15A
EF3 200A DC ESS BUS 1 BC 1 200A EF4 225A
DCTC 120A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
AC CIRCUIT BREAKERS
RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
EC 1 120A DC ESS BUS3 EF1 EF2 150A 150A
GLC 2 120A
BTC 2 120A
AC BUS 1
CB26 25A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
BATT 2
TO APU START DC EXT PWR
NAME
SOURCE
CB23
SPARE
AC ESS BUS NONE
CB24
FLAP 1 AC FEED
CB25 CB26
FWD FAN FEED TRU ESS FEED
AC ESS BUS 5 A AC 3 PHASE AC ESS BUS 25 A AC 3 PHASE
CB27
SPARE
AC ESS BUS NONE
CB28
AC FUEL PUMP 2A
AC ESS BUS 7.5 A AC 3 PHASE
CB29
SLAT 2 AC FEED
AC ESS BUS 15 A AC 3 PHASE
CB30
FAN FEED
AC ESS BUS 5 A AC 3 PHASE
NAME
SOURCE
CB RATING
CB21
EXCITER 1A
STANDBY AC BUS
CB22
EXCITER 2A
STANDBY AC BUS
5 A AC 1 PHASE 5 A AC 1 PHASE
AC CIRCUIT BREAKERS
CB RATING
15 A AC 3 PHASE
Essential Integrated Control Center EICC Take precautions to ensure proper power and Electro Static Discharge (ESD) safety considerations while performing maintenance on the open ICCs. Electrical power must be disconnected an ESD wrist strap must be worn since there are static sensitive modules installed in the ICCs.
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24-60 DC Power Distribution System Introduction Direct Current (DC) power distribution routes primary DC power to the aircraft through subassemblies located in the Left Integrated Control Center (LICC), Right Integrated Control Center (RICC), Emergency Integrated Control Center (EICC) and Auxiliary Integrated Control Center (AICC). The two main DC buses (DC BUS 1; DC BUS 2) can receive primary DC power from two main DC channels coming from Transformer Rectifier Units (TRU 1) or 2. The three Essential DC buses (DC ESS BUS 1, DC ESS BUS 2, DC ESS BUS 3) receive power from either of the main DC buses, the ESS DC channel out of the ESS TRU, the external DC power channel or either of the two battery channels. Both main DC buses and the three ESS DC buses supply power to batteries 1 and 2 via the battery buses. These buses also supply the AC inverter. The left main DC bus supplies power to the DC Ground Service Bus (DC GND SVC).
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Page 1
Figure 1: Schematic
RAT GEN
GLC 1 120A
EICC RLC 60A
APU GEN
EPAC 120A
CB26 25A
AC GND SVC
STANDBY AC BUS TRU ESS 300A
DC GND SVC
TRU EC 400A
CB4 35A TRU 1 300A
TRU2 300A
TRU 1C 400A DC BUS 1
ETC1 120 A
CB17 15A
DCTC 120A
EF3 200A DC ESS BUS 1 BC 1 200A EF4 225A
TRU 2C 400A DC BUS 2 EC 2 120A RF2 ETC 2 200A 120A DC ESS BUS2 BC2 200A RF1 150A
AICC
AF1 225A HOTBATT BUS 2 ABC 400A
HOTBATT BUS 1
AC INVERTER 250VA DC
Issue: June06 Revision: 00
RICC
AC BUS 2 CB1 CB29 50A 35A
GSTC 60A
EC 1 120A DC ESS BUS3 EF1 EF2 150A 150A
GLC 2 120A
BTC 2 120A
AC BUS 1 CB3 50A
IDG 2
ALC 120A
LICC
BTC 1 120A
AETC 60A
AC ESS BUS STBYC 10A
AC EXT PWR
IDG 1
BATT 1
FOR TRAINING ONLY Reproduction Prohibited
ISOLATED RICC AF2 ASC 300A 400A
EPDC 400A
TO APU START DC EXT PWR
BATT 2
Chapter 24-60
Page 2
170/190 MAINTENANCE TRAINING MANUAL
Intergrated Control Centers and SPDAs All four Integrated Control Centers have circuit breakers that provide DC power control and protection to a selection of aircraft electrical loads. The two SPDAs are an integral part of the Electrical Power Generation and Distribution System (EPGDS). The SPDAs provide control and monitoring of the primary DC distribution and of DC secondary loads. Other DC system control is accessible using manual switches on the electric system cockpit control panel. The two main DC buses, three essential DC buses and one ground service bus distribute power to various aircraft electrical loads either through DC circuit breakers or through Secondary Power Distribution Assembly One or Two.
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Chapter 24-60
Page 3
Figure 2: Intergrated Control Centers and SPDAs
SPDA 2
SPDA 1
EIC EICC
RICC
ELEC AC SYSTEM IDG 1 CONTROL DISC
DISC
IDG 2 CONTROL
AC BUS 1 GPU AVAIL
AC BUS 2 AC BUS TIES AUTO
1 OPEN
APU GEN
2 OPEN
IN USE
ELEC DC SYSTEM TRU 1
DC BUS 1
TRU ESS
TRU 2
ESS DC BUS
DC BUS 1
BATT 1 ON OFF
BATT 2 DC BUS TIES
AUTO OFF
LICC AICC
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170/190 MAINTENANCE TRAINING MANUAL
24-61 Secondary Power Distribution Assemblies Introduction Secondary DC power is distributed by LRUs located in the Secondary Power Distribution Assembly (SPDA 1 and SPDA 2). Secondary Power Distribution Assembly, (SPDA 1) is located in the forward avionics bay and SPDA 2 is located in the mid avionics bay. In the DC distribution system, SPDA 1 provides secondary power control and monitoring of three Integrated Control Centres (LICC, RICC and EICC) and SPDA 2 provides secondary power control and monitoring of all four ICCs. SPDA 1 consists of twenty slots for modules; SPDA 2 has twenty-six modules that are individually replaceable. Some of these modules assist in DC power distribution. An EICAS message "SPDA FAIL" indicates a total failure of any module in SPDA 1 or SPDA 2, or a loss of communication with the other SPDA. This indication is shown as an advisory message.
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Chapter 24-61
Page 1
Figure 1: SPDA Locations
SPDA 2
SPDA 1
EI EICC
RICC
ELEC AC SYSTEM IDG 1 CONTROL DISC
DISC
IDG 2 CONTROL
AC BUS 1 GPU AVAIL
AC BUS 2 AC BUS TIES AUTO
1 OPEN
APU GEN
2 OPEN
IN USE
ELEC DC SYSTEM TRU 1
DC BUS 1
TRU ESS
ESS DC BUS
TRU 2
DC BUS 1
BATT 1 ON OFF
BATT 2 DC BUS TIES
AUTO OFF
LICC AICC
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170/190 MAINTENANCE TRAINING MANUAL
SPDA Function • Integrate electric power distribution/control with aircraft utility system control • Electrical power received from four independent DC buses SPDA 2 SPDA 1 DC BUS 1 DC BUS 2 DC ESS BUS 1 DC ESS BUS 1 DC ESS BUS 2 DC ESS BUS 2 DC GND SVC DC GND SVC
• Monitoring and distribution control of aircraft utility system Air Management Oxygen Electrical Power Engine Ignition Fuel Engine Starting Hydraulics APU Anti Ice Fire Extinguishing Lighting Water
• Distribution control via Solid State Power Controllers (SSPCs) • Trip characteristics match traditional circuit breakers No moving parts.
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Page 3
Issue: June06 Revision: 00 FOR TRAINING ONLY Reproduction Prohibited 26
POWER SUPPLY
AMS PROCESSOR
7,5
CB45
DC BUS 2
25
5
AMS LOW LEVEL I/O
AMS - MTR DRIVE
AMS LOW LEVEL I/O
25
CB17
CB34
24
23
22
AMS LOW LEVEL I/O
AMS - MTR DRIVE
AMS LOW LEVEL I/O
AMS PROCESSOR
ARINC 429
FILLER
FILLER
AC POWER
AC POWER
5
CB50
AC BUS 2
21
20
19
18
17
16
DISCRETE I/O MICRO/COMMS
5
25
CB14
25
RICC
15
DISCRETE I/O ASCB COMMS
CB07
25
CB12
25
CB11
7,5
DC ESS 2
CB54
CB32
5
CB39
5
CB38
7,5
CB34
25
CB16
25
CB31
25
CB50
LICC
14
13
12
11
10
9
MICRO/COMMS
DC POWER
DC POWER
DC POWER
DC POWER
EICC
8
7
6
5
MAU 3 4
APU FADEC
CB08
5
DC POWER
AC GND SVC
3
CB40
5
7,5
7,5
CB33
15
CB24
CB42
DC POWER
CB10
7,5
7,5
CB48
CB41
DC ESS 3
5
DC ESS 1
CB13
5
CB01
25
5
25
25
25
7,5
DC GND SVC
POWER SUPPLY
POWER SUPPLY
DISCRETE I/O
ARINC 429
ANALOG
AC POWER
FILLER
FILLER
DC POWER
CB06
CB49
25
CB10
CB09
CB30
CB51
DC BUS 1
2
1
20
19
18
17
16
15
DISCRETE I/O MICRO/COMMS
5
CB14
25
CB07
CB06
RICC
14
13
DISCRETE I/O ASCB COMMS
25
CB05
7,5
CB04
EICC
12
11
10
MICRO/COMMS
DC POWER
DC POWER
DC POWER
DC POWER
DC POWER
DC POWER
POWER SUPPLY
LICC
9
8
7
6
5
4
3
2
1
EMBRAER 170 / 190
Figure 2: SPDA 1 Power Supplies GPU
CB44 EPM
GCU 1
DC ESS 2
DC BUS 2
FOR REFERENCE ONLY SPDA 2 AMS Ch B
MCDU 1
SPDA 1 DC BUS 1
AC BUS 1 DC GND SVC
EPM
5
GCU 2
5
AGCU
DC ESS 3 DC ESS 1
FOR REFERENCE ONLY SPDA 1 ARINC 429
MCDU 2
SPDA 2
Chapter 24-61 Page 4
170/190 MAINTENANCE TRAINING MANUAL
SPDA 1 Location SPDA 1 is located in the temperature and pressure controlled forward EBay. Access to this equipment can be accomplished through the forward EBay floor access hatch, which is located in front of the nose gear. SPDA 1 mounts to the aircraft rack using 8 screws (size 10-32). The screws are installed from inside the chassis into captive fasteners in the aircraft rack. The electrical connectors for SPDA 1 are on the back panel. Cooling air is provided from the aircraft Air Management System (AMS) and drawn through SPDA 1 to ensure adequate heat dissipation for the SPDA 1 modules (20). A Cooling Plenum with two ports is located near the top of SPDA 1. Air inlet holes are on the bottom and on the top. SPDA 1 requires 150 cubic feet per minute of cooling airflow, requiring 1.28 inches of water pressure (under normal conditions of two fans operating at sea level and 40oC). The position of the Identification Plate for each LRM is on the Insert/Extractor and Stiffener. SPDA 1 has 20 modules. SPDA1 is not considered a line replaceable unit (LRU), but each of the 20 modules is considered to be a line replaceable module (LRM). SPDA 1 and its LRMs should be handled using static discharge prevention equipment and practices. Dimensions and weight for SPDA 1 (excluding Cooling Plenum) are as follows: • Height = 8.7 in (222 mm) • Depth = 11.8 in (299 mm) • Width = 22.2 in (565 mm) • Weight = 58.7 lbs (26.6 kg) The weight of the Cooling Plenum for SPDA 1 is 3.7 lbs (1.7 kg).
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Figure 2: SPDA 1 Location
Fwd
SPDA 1 EICC
Forward E-bay RAT GCU
SPDA1 EICC
Battery 1 Static Inverter Forward E-Bay
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SPDA 2 Location SPDA 2 is located in the temperature and pressure controlled Mid EBay. Access to this equipment can be accomplished through the Mid E-Bay floor access hatch, which is located on the aircraft port side, behind the left wing. SPDA 2 mounts to the aircraft rack using 8 screws (size 10-32). The screws are installed from inside the chassis into captive fasteners in the aircraft rack. The electrical connectors for SPDA 2 are on the back panel. Cooling air is provided from the aircraft AMS to ensure adequate heat dissipation for the SPDA 2 modules (26). A Cooling Plenum with two ports is located above and near the front of SPDA 2. Air inlet holes are on the bottom and on top. SPDA 2 requires 200 cubic feet per minute of cooling airflow, requiring 1.16 inches of water pressure (under normal conditions of two fans operating at sea level and 40oC). The position of the Identification Plate for the chassis is on the left side. The position of the Identification Plate for each LRM is on the Insert/Extractor and Stiffener. SPDA 2 has 26 modules (including eight for the AMS). SPDA 2 is not considered a line replaceable unit (LRU), but each of the 26 modules is considered to be a line replaceable module (LRM). SPDA 2 and its LRMs should be handled using static discharge prevention equipment and practices. Dimensions and weight for SPDA 2 (excluding Cooling Plenum) are as follows: • Height = 8.7 in (222 mm) • Depth = 11.8 in (299 mm) • Width = 28.5 in (724 mm) • Weight = 66.1 lbs (30.0 kg) The weight of the Cooling Plenum for SPDA 2 is 4.6 lbs (2.1 kg).
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Page 7
Figure 3: SPDA 2 Location
SPDA 2
AMS Cooling Ducts
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Mid E-bay
Fwd Mid E-bay
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170/190 MAINTENANCE TRAINING MANUAL
SPDA Modular Composition and Power Sources DC Power I/O Modules provide the following: - Four 2.5 to 15 Amp Outputs - Eight 2.5 to 7.5 Amp Outputs - One 7.5 Amp Isolated FET - Four Ground/Open Inputs - Four 28V/Open Inputs All outputs are individually programmable with respect to current rating and default state. AC Power I/O Modules provide the following: - Nine 2.5 to 7.5 Amp Outputs - Four Ground/Open Inputs - Four 28V/Open Inputs All outputs are individually programmable with respect to current rating and default state. Power outputs can be used separately or configured in groups of three with the same current rating and default state for three phase loads. Discrete I/O Modules provide the following: - Twelve 250mA 28V/Open Outputs - Twelve 250mA Ground/Open Outputs - Twelve 28V/Open Inputs - Twelve Ground/Open Inputs All outputs above are individually programmable with respect to default states. Analog I/O Module provides the following: - Six 115 VAC voltage monitors - One 500A high side current monitor - Two 500A low side current monitors - One RTD temperature monitor - Four 28VDC voltage monitors - One combined 28VDC voltage and ripple monitor - Two voltage ripple monitors for 28VDC supplies
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- Two counter inputs for speed/frequency measurements ARINC 429 Communication Modules provide the following: - Eight ARINC-429 Receivers - Four ARINC-429 Transmitters Receiver channels can be independently programmed to receive 12.5kHz or 100kHz ARINC data. Label filtering is provided to ignore data not intended for use by SPDAs. Transmitter channels are independently programmable for 12.5kHz or 100kHz operation. ASCB Modules provide the following: - Transmit function on two ASCB-D busses - Receive function on three ASCB-D busses - Transmit and Receive on one LAN - Discrete inputs for module ID assignment Microprocessor Modules provide the following: - Cross-channel SPDA communication interface (CAN) - Master chassis Identification - Application identification pin strapping Power Supply Modules provide the following: - Dedicated Power Supply for each Microprocessor Module - Distributed power supply for other modules - Redundant 28VDC inputs Filler Modules are utilized where provision for another module type has been made but is not utilized. They protect the back plane connector and installation guides from FOD. A cover plate is used to cover the aircraft side of the EPXB connector to satisfy EMI and environmental requirements.
Height – 8.7 in (222 mm) Depth – 11.8 in (299 mm) Width – 22.2 in (565 mm) Weight – 58.7 lbs (26.6 kg)
SPDA Power Supply Module
Wedge Lock Issue: June06 Revision: 00
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170/190 MAINTENANCE TRAINING MANUAL
SPDA 2 Line replaceable Modules (LRMs) SPDA 2 contains twenty-four LRMs and two filler modules for the two spare slots. SPDA 2 utilizes a 26 slot chassis, which houses not only the SPDA components but also the Air Management System (AMS) control modules. Ducted AMS cooling air is utilized for SPDA cooling. The SPDA LRMs fall within 3 categories: - Power supply - Microprocessor - Input/Output (I/O) Power Supply LRMs are utilized in pairs to provide redundant control power for all of the other LRMs in a SPDA chassis. They are located on each end of the chassis for the best thermal environment. Microprocessor LRMs are also utilized in pairs for redundancy. They implement all of the control laws of the utility management system coordinating SPDA operation and interface to the aircraft avionics system (flight deck. Since they are the second highest power dissipating LRMs after the power supply LRMs, they are also placed where they can obtain the best thermal environment. They are not adjacent to the power supply LRMs. There are three slots between the two microprocessor LRMs. The ASCB LRM is located between the two microprocessor LRMs since it must share the common PCI bus. Since the ASCB LRM is also high power dissipating, it is installed in the middle of these three slots so that it is not directly next to either microprocessor LRM. The Input/Output (I/O) LRMs fall within the following five categories: - Communication (ASCB and ARINC 429) - Discrete I/O - Analog I/O - DC Power I/O - AC Power I/O Communication LRMs are used to communicate with the avionics computers and can be used as sub system interfaces to display units, local control panels or other sub system controllers.
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Discrete I/O LRMs interface with aircraft switches, sensors and low current loads such as indicator lights. Analog I/O LRMs interface with aircraft sensors and control elements. DC and AC Power I/O LRMs provide the dual function of providing a circuit breaker function and output control.
FOR TRAINING ONLY Reproduction Prohibited
Chapter 24-61
Page 13
Figure 6: SPDA 2 (module location)
ASCB Comms Micro & Comms
Spares AMS AMS Motor Power Micro Drive Supply
Micro & Comms
ESD Jack
Slot 1
Slot 26
Power Supply Issue: June06 Revision: 00
DC Power
Discrete I/O
AC Power
AMS I/O ARINC 429
FOR TRAINING ONLY Reproduction Prohibited
AMS Micro Chapter 24-61
Page 14
170/190 MAINTENANCE TRAINING MANUAL
Notes:
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Chapter 24-61
Page 15
Figure 7: SPDA Power Supply Module
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170/190 MAINTENANCE TRAINING MANUAL
SSPCs SSPC Functional Separation SSPC circuit breaker and switch functions utilize independent control logic. After CB ratings are programmed at initial power up, the CB command function operates independently of the switch command function. Three Phase AC SSPCs Unless otherwise specified, all three phase AC SSPCs turn ON, OFF, and trip on over current within 10 ms of each other. DC SSPCs used in Parallel DC SSPCs within the same SPDA, which are used as Diode “or’d” sources controlled by common logic, turn OFF within 10 ms of each unless staggered for BIT, but they trip independently in the event of over current. SSPC Initialization When power is first applied to an SPDA, all SSPCs initially have the CB and switch functions set to OFF and verified to be OFF. SSPC switch ON are independent of the state of the associated power bus unless dictated otherwise by control logic such that when the power feed or bus subsequently becomes available, all “on” outputs on that module simultaneously become live. SSPCs do not turn ON unless the following requirements are satisfied: - The switch state is ON - The CB function is IN - Programmed CB rating has been verified
When the operating schedule is enabled and running, all unused SSPCs are commanded OFF. Discrete Outputs Not Used All unused discrete outputs are configured with the default state OFF. Discrete Outputs are commanded OFF.
SSPCs Used as Power Feeds SSPCs used as power distribution feeds only and which have no associated control logic, are switched ON once the defined current rating has been programmed and verified. SSPCs Not Used All unused SSPCs are configured as follows: - Circuit Breaker rating is 2.5A - The default state is OFF Issue: June06 Revision: 00
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Chapter 24-61
Page 17
Figure 8: SPDA 1 (module partially removed)
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170/190 MAINTENANCE TRAINING MANUAL
CAS Advisory Message REMOTE CB TRIP The EPGDS control units (GCU1, GCU2, AGCU, SPDA1, and SPDA2) generate the CAS Advisory Message REMOTE CB TRIP when one of the following circuit breakers (CBs), fuses, or Solid State Power Controllers (SSPCs) are tripped or blown: - GCU1 - LICC thermal circuit breaker (CB) detected as tripped - LICC fuse detected as blown - GCU2 - RICC thermal circuit breaker (CB) detected in tripped state - RICC fuse detected as blown - AGCU - AICC thermal circuit breaker (CB) detected as tripped - AICC fuse detected as blown - SPDA1 - EICC thermal circuit breaker (CB) detected in tripped state - EICC fuse detected as blown - SPDA1 SSPC detected as tripped - SPDA2 - SPDA2 SSPC detected as tripped The message is inhibited by CAS (MAU) during takeoff, landing, and while in flight.
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Chapter 24-61
Page 19
Figure 9: MCDU CB Control
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Page 20
170/190 MAINTENANCE TRAINING MANUAL
Notes:
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Chapter 24-61
Page 21
Figure 10: MFD Electrical Synoptic Page
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170/190 MAINTENANCE TRAINING MANUAL
MCDU CB Control Page • Select CB button on MCDU • CB pages show all CBs on aircraft (virtual and mechanical CBs) • Virtual CBs (inside SPDAs) can be pushed in or pulled out by selection on MCDUs • Mechanical CBs have indications if in or out and location labels (LICC, RICC, EICC) • MCDU 1 controls some CBs MCDU 2 controls other CBs
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Chapter 24-61
Page 23
Figure 11: MCDU CB Control Page
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Page 24
170/190 MAINTENANCE TRAINING MANUAL
24-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ 42-00 AC External Power C ¦ 1 ¦ System ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦
1) AC GPU AVAIL/IN USE Pushbutton Lights
C ¦ 4 ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
2) AC GPU C ¦ 1 AVAILABLE ¦ Light on ¦ Flight ¦ Attendant ¦ Ground Service ¦ Panel ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
3) AC GPU IN USE C ¦ 1 Light on ¦ Flight ¦ Attendant ¦ Ground Service ¦ Panel ¦
¦ 0 ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
¦ 52-03 In-Flight C ¦ 1 ¦ *** Entertainment ¦ ¦ System (IFE) Auto ¦ ¦ Shutdown ¦
¦ 0 ¦ ¦ ¦
¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 24-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 24 ELECTRICAL POWER ¦ ¦ ¦
¦ 00-05 Batteries 1 and 2 C ¦ 4 ¦ Voltage ¦ ¦ Indication on MFD ¦ ¦ Status Page ¦
¦ 2 ¦ ¦ ¦
¦ One indication per battery may be ¦ inoperative. ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦
C ¦ 4 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ affected battery voltage is ¦ available on MFD Electrical Page.
¦ ¦ ¦
¦ 22-01 APU Generator ¦
C ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided APU ¦ generator remains selected off.
¦ ¦
¦ 36-10 Batteries 1 and 2 C ¦ 4 ¦ Temperature ¦ ¦ Sensors ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 2 ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 41-00 DC External Power D ¦ 1 ¦ *** System ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
1) DC GPU AVAIL/IN USE Pushbutton Lights
D ¦ 2 ¦ ¦ ¦
Any indication on MFD Electrical Page may be inoperative.
| ¦ | ¦ | ¦
One sensor per battery may be inoperative provided at least one temperature of associated battery on Electrical Synoptic Display (MFD Electrical Page) is verified to operate normally before each flight.
31-21 Clock system General The Clock is located in the cockpit instrument panel. It is used to display the Universal Time Coordination (UTC), elapsed time, and chronograph functions.
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Chapter 31-21
Page 1
Figure 1: Clock system
CHR
RST
CHR
SEC
MIN
UTC
GPS INT
SET
DATE
SET HR/MO
MIN/DY
AUTO RST
HR
CHR
RST
SEC/Y
ET
MIN
CHR
SEC
MIN
UTC
GPS INT
SET DATE
SET HR/MO
MIN/DY
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SEC/Y
AUTO RST
HR
ET
MIN
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170/190 MAINTENANCE TRAINING MANUAL
The clock display The Clock interfaces with the Generic Input/Output Module 2 in MAU 2 via an ARINC 429 Bus. The Clock receives GPS data and a Weight On Wheels signal, but provides UTC and date. Power is supplied from the 28 VDC Essential Bus 1. The front panel consist of high contrast transflective LCDs with blue-white segments on black background and control buttons. There are four digits for the chronograph functions, six digits for Universal Time Coordination, and four digits for elapsed time. The LCDs are back lighted and automatically dimmed using a photo cell. Two push buttons control the chronometer: Chronometer (CHR) and Reset (RST). A three-position knob selects Global Positioning System time (GPS), internal time (INT) or sets Universal Time Coordination and date functions (SET) for the UTC display. The DATE/SET push and turn button allows you to set the date. A two-position mode selector for the elapsed time is used to select the automatic mode (AUTO), or reset function (RST).
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Figure 2: Clock Display High contrast transflective LCDs Control buttons: Photo cell
Reset (RST)
CHR
RST
UTC and Date
Chronometer (CHR) SEC
MIN
Generic Input/Output Module 2 MAU #2
CHR
UTC ARINC 429
GPS INT
Internal Time (INT)
SET
DATE
Global Positioning System time (GPS)
SET HR/MO
MIN/DY
SEC/Y
Set for UTC display (SET)
AUTO RST
GPS data Weight On Wheels signal
Automatic mode (AUTO) Reset (RST) HR
ET
MIN
DATE/SET push and turn button 28 VDC Essential Bus 1
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170/190 MAINTENANCE TRAINING MANUAL
The clock functions UTC is indicated from 0 to 23 hours 59 Minutes and 59 Seconds. A fixed colon separates the hours from the minutes and comes on only when the function is initialized. With the three-position selector set to INT, the clock displays its internally computed time. When set to GPS, the clock synchronizes and displays GPS time. Note: During initialization and set to GPS the clock displays its internal time. When the clock is synchronizing with the GPS, when no signal is present or when false or invalid data is received, the clock displays its internal time. It displays all dashes if the internal time is not set. Pressing the DATE button displays the DAY/MONTH and YEAR in place of the UTC, but the colon is not displayed. Leap years are taken into account. Elapsed time counting is possible from 0 to 99 hours and 59 Minutes. A fixed colon separates the hours from the minutes and comes on only in the RUN mode. The display is blanked in the non operative mode. In the AUTO mode, the elapsed time counter starts only if there is no Weight On Wheels (WOW is false). With Weight On Wheels (WOW is true) and selected to the RESET mode, the counter is zeroed and blanked. Chronometer time is displayed from 0 to 99 Minutes and 59 Seconds. Pushing the RST push button blanks the display. Pushing the CHR push button enables the display and starts the chronometer from zero. Pushing it a second time stops it. The display is blanked in the non-operative mode.
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Figure 3: The clock functions UTC time
CHR
RST
The clock synchronizes and displays GPS time
CHR
UTC
GPS INT
SET
The clock displays its internally computed time
SET MIN/DY
SEC/Y
AUTO RST
HR
ET
CHR
SEC
MIN
UTC
HR/MO
CHR
RST
SEC
MIN
DATE
Displays all dashes if the internal time is not set
GPS INT
SET
DATE
SET HR/MO
MIN/DY
AUTO RST
HR
MIN
SEC/Y
ET
MIN
Note: During initialization and set to GPS the clock displays its internal time. When the clock is synchronizing with the GPS, when no signal is present or when false or invalid data is received, the clock displays its internal time. When Date button is pressed: Date is displayed instead of UTC
Elapsed time counting: From 00:00 to 99:59
Blanks the display Starts when: No Weight On Wheels
CHR
RST
CHR
UTC
HR/MO
MIN/DY
SEC/Y
AUTO RST
ET
GPS INT
INT SET
HR
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UTC
GPS
SET
DATE
CHR
MIN
SET
DATE
SET HR/MO
(WOW is false)
MIN/DY
SEC/Y
Stops when: Weight On Wheels (WOW is true)
AUTO
HR
Starts the chronometer from zero
ET
CHR
RST
SEC
MIN
SEC
MIN
CHR
RST
Chronometer time from: 0 to 99 Minutes and 59 Seconds
FOR TRAINING ONLY Reproduction Prohibited
SEC
MIN
UTC
GPS INT
SET
DATE
SET HR/MO
MIN/DY
SEC/Y
AUTO RST
HR
MIN
CHR
ET
MIN
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31-60 Central Display System Introduction The Electronic Display System (EDS) presents primary flight, navigation, and system information to the flight crew. The EDS consists of: • • • • • • •
5 display units, a guidance panel, 2 Cursor Control Devices (CCD), 2 Multi Function Control Display Units (MCDU), EICAS declutter panel, 2 reversionary panels, MAU hardware including Control input/output modules, and an EDS application software on dedicated processor modules.
The five display units are located on the main instrument panel. There are two Primary Flight Displays (PFD), two Multi function Displays (MFD), and an Engine Instrument and Crew Alerting System (EICAS) display. The display units are identical. In case of failure of one display, an automatic logic transfer will allow its information to be presented in the remaining units. The Multi Function Control Display Units (MCDU) may be used as a back up for the main panel displays.
Figure 1: Electronic Display System (EDS)
Electronic Display System (EDS)
5 display units Guidance panel 2 Cursor Control Devices (CCD) 2 Multi Function Control Display Units (MCDU) EICAS declutter panel reversionary panels MAU hardware including Control input/output modules EDS application software on dedicated processor modules
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EFD Power Supply All of the components require 28 VDC power. The pilot's PFD, the copilot's MFD, the copilot's CCD, the pilot MCDU and copilot's display controller are powered by DC Bus 1. The copilot's PFD is powered by DC Bus 2. The pilot's MFD and the pilot´s CCD are powered by the DC Ess Bus 2. The EICAS is powered by the DC Ess Bus 1. The pilot's display controller is powered by DC Ess Bus 2 via SPDA 2.
Figure 2: EDS Power Supply
V DC)
DC Bus 1 lot's lot s PFD opilot's MFD opilot's CCD lot's MCDU opilot's Disply Contoler
us 2 opilot s PFD opilot's
Ess Bus 2 ot's ot s MFD ot's CCD ot's Display Controller (via SPDA 2)
Ess Bus 1 CAS
last update: Nov06
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Electronic Display System (EDS) The Electronic Display System (EDS) has four basic functions: graphics generation, monitor warning, control abstraction layer, and utility. The graphics generation function generates flight information such as primary flight displays, navigation displays, crew alerting messages, engine systems and weather information, cockpit annunciation, radio tuning information and terrain awareness and warning information. The monitor warning function integrates the aural warning and crew alerting functions. The monitor warning receives warning, caution, advisory and status signals from various systems in the aircraft and will perform the same computations on the identical signals independently. The control abstract layer function resides in the control I/O modules, processor modules and display units in order to perform radio tuning pages, avionics setup page, avionics test pages and thrust management pages for the Multi Function Control Display Units (MCDUs). It also provides centralized control for on screen cursor and knob data from the CCD. The utility function is a software that provides ASCB command data to the I/O modules for control of the aircraft systems.
Figure 3: Table EDS Functions
Graphics generation
Monitor warning
Control abstraction layer
Utility
Primary flight displays
Warning signals
Radio tuning pages
Navigation displays
Caution signals
Avionics setup page
Software that provides ASCB command data to the I/O modules for control of the aircraft systems.
Crew alerting messages
Advisory signals
Avionics test pages
Engine systems and weather information
Status signals
Thrust management pages
Cockpit annunciation
Centralized control for on-screen cursor and knob data from the CCD
Radio tuning information
Terrain awareness and warning information
last update: Nov06
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EDS Abnormal Function The Electronic Display System (EDS) abnormal operation is the automatic and forced Display Units (DU) reversion. The five display units are numbered 1 to 5 from left to right. Only DU 2 and DU 4 can be reverted. DU 1 and DU 5 are always operated as PFDs, and DU 3 is always an EICAS. Each pilot can manually select the desired display on the MFD screen via the display reversionary panel. The four switch positions are PFD, AUTO, MFD, and EICAS. Setting the switch to a position other than AUTO forces that selection onto the MFD. In this case the “failed/ reverted from” display unit is shutdown and the display will be blank. When the switch is set to the AUTO position, the system is in automatic reversion (with manual override capability) and it will detect many types of display unit failures. The reversion system will reconfigure the remaining DUs to the configurations shown.
Guidance Panel The guidance panel houses two display controllers and a flight guidance panel. The display controller enables the selection of primary flight display HSI formats, navigation sources, weather display and bearing pointer selection. The flight guidance portion allows selection of autopilot and yaw damper engagement functions, flight director mode engagement and selection of display data source for the flight director.
Autopilot and yaw damper engagement Flight director mode engagement Display data source for the flight director
last update: Nov06
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The Display Controller The display controller has following controls: The combined BARO SET/ STANDARD/UNITS knob has three functions: Turning the BARO knob sets barometric altimeter correction, pushing the PUSH STD button sets baro correction to standard, and selecting the IN/HPA switch selects baro correction format to inches of mercury or to Hecto Pascals. The combined RA/BARO-MINIMUM/MODE knob has three functions: The outer knob selects the source, either barometric or from the radar altimeter system. The inner knob when turned clockwise increases the Decision Height (DH). The HSI button toggles between full HSI and partial compass (arc) displays. The WX button selects weather data for display. The FMS button selects FMS as the primary navigation source and toggles between FMS sources. The BRG buttons toggle between VOR, ADF and FMS bearing sources; the upper button - circle - is for system 1, and the lower button - diamond - is for system 2. The PREV button selects navigation preview. When FMS is the selected navigation source, it is used to display VOR/LOC lateral and vertical deviation and distance on the HSI. The V/L button toggles between available short-range navigation sources. The FPR button turns the flight path reference line on the ADI on and off.
Figure 6: The Display Controller
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31-61 Displays Primary Flight Display The Primary Flight Display (PFD) is the flight crew’s primary instrument. It provides display of aviation and navigation information, as well as backup radio tuning. The PFD is divided into sections, each one presenting one group of information such as information regarding attitude, heading, flight modes, and tuning COM and NAV. On the guidance panel the display controller portion allows the selection of primary flight display HSI formats, navigation sources, weather display, and bearing pointer selection. Certain PFD internal failures will result in a large red X covering the PFD screen. In case of mismatched information between the two PFDs, no information at all will be presented. In the event of a display failure, information may be presented in the MFD by appropriately setting the reversionary panel.
Display Units The function of each DU is related to its position in the cockpit panel. The DU is an 8” x 10” LCD (Liquid- Crystal Display) that has an autonomous processing capability and is directly connected to the ASCB and LAN (Local Area Network). The DUs contain the components that follow: • LCD • NIC (Network Interface Controller) The NIC controls the data between the DU, ASCB and LAN. • Processor The processor controls the I/O (Input/ Output), graphics and display functions in the DU. • I/O The I/O supplies the necessary interfaces to the ARINC-429 buses and discretes. • Backplane bus • Cooling fan The DU positions in a normal configuration are as follows: • PFD (DUs No. 1 and No. 5) • MFD (DUs No. 2 and No. 4) • EICAS (DU No. 3)
Figure 2: Display Units
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Notes:
Figure 3: Displays
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Multi Function Display The Multi Function Display (MFD) presents navigation information, shows aircraft's system parameters and status, can provide a pilot's checklist, and enables maintenance personnel to access maintenance messages. The MFD display consists of control and information areas. The menu soft keys, on the top and bottom of the screen, are used to select formats and control various systems. The selected information is displayed on the split up screen. The Cursor Control Device (CCD) in the centre pedestal is used to control the menu soft keys. Once a soft key is pushed, a pull down menu is opened. Check boxes in the menu are used to select and deselect each function. For navigation you can choose between the map and plan format. Weather radar, FMS, TCAS, and various other information can be shown or alternatively switched on and off.
Figure 4: Multi Function Display (MFD)
Multi Function Display (MFD) vigation information MAP
Plan
Systems
Fuel
ircraft's system parameters and status an provide a pilot's checklist nables maintenance personnel to access aintenance messages
TCAS
last update: Nov06
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WX
Checklist
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Multi Function Display (continued) The systems menu has 8 sub pages available. 7 pages are synoptic displays for monitoring the respective aircraft system and for fault isolation. These are the pages for: Status, Flight Controls, Hydraulics, Fuel, Electrical, Environmental Control System (ECS), and Anti-Ice. Selecting the maintenance page permits to access maintenance messages. The maintenance page is available only on right MFD and on ground and is only for maintenance personnel. Pushing the TCAS soft key opens the TCAS controller. Actuating the weather button pops up the weather radar virtual controller. Via the checklist soft key you access the electronic checklist, which allows to select among normal, abnormal or emergency procedures indexes.
Figure 5: Multi Function Display cont.
MAP
Plan
Systems
Fuel
Status Ctrl Flight Ctrl Hudralics Fuel Electrical ECS Anti - Ice Maintenance
The maintenance page is available only on right MFD and on ground and is only for maintenance personnel.
TCAS
last update: Nov06
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WX
Checklist
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System Status Page Engine Bleed Pressure Digital Readouts The Engine Bleed Pressure Digital Readouts have a label of PSI below the digital readout. The Bleed Pressure is provided by the Air Management System and indicates the current Bleed Air Pressure from the respective engine. Flight Status The Flight Status consists of the Flight Number, Flight Time, Total Air Temperature, Static Air Temperature and Gross Weight. The top left corner of the MFD Status Page shows the Flight Status. Flight Number The Flight Number display is shown in an alphanumeric digital readout with the label FLIGHT to the left. The Flight Number data is provided by the coupled - side (priority) FMS. Crew Oxygen Display Crew Oxygen Pressure display consists of the oxygen pressure information required for crew awareness. The analog scale displayed is partitioned in to three segments. The lower 50 % of the scale is represent by amber bar indication low oxygen pressure. The next 25 % of the scale represented by a cyan bar, indicates oxygen sufficient for a crew of two. The final top 25 % of the scale, represented by the white bar, indicates oxygen sufficient for a crew of three. Amber tic marks will be displayed at the scale‘s bottom and mid - point. At 75 % of scale‘s length, a cyan tic mark will be displayed while a white tic mark will be displayed at the top of the scale. The scale‘s pointer color is directly tied to the color of the digital readout. Engine 1 and 2 Oil Scale and Pointers The Engine Oil Level Scale is in an analog display of the engine oil level in quarts. Each engine has its own scale. The scale is divided into two segments. The bottom segment is the amber region ranging from 0 to the minimum oil level sent by the correspondent FADEC. The upper segment is displayed in white and is the normal region for values greater than the minimum oil level. There is a tic mark at the minimum oil level to seperate the
two segments. The scale is a vertical line with tic marks at each end, a label ENG OIL LEVEL displayed at the top, and a units label QT at the bottom. Each scale has its Engine Oil Level Pointer. Each pointer is positioned in the external side of each scale. The left Scale corresponds to engine 1 and the right scale is for engine 2. In the normal range the pointers are green. In the amber region, the pointers are amber. Brakes System Emergency Accumulator Pressure Scales and Pointers The Brakes System Accumulator Pressure display has two scales, one for each brake system. The Brakes System Accumulator Pressure Scales are labeled SYS1 and SYS2, respectively. The scales are divided into two segments. The top of the scale is the normal region, with a range of 1900 psi to 4000 psi for the 170 aircrafts and 1200 psi to 4000 psi for the 190 aircrafts. The bottom portion of the scale is the amber region, ranging from 0 to 1200 psi for the 190 aircrafts, and 0 to 1900 for the 170 aircraft. there is a tic mark at 1200 psi (1900 psi) to seperate the segments, and tic marks at each end of the scale. The amber segment indicates that the accumulator pressure is below half of the total available pressure and is enough only for three brake applications. Each scale has a Brakes System Accumulator Pressure Pointer to indicate the accumulator pressure. In the normal range, the pointer is green. In the amber range, the pointer is displayed as solid amber. Brakes System Temperature Scales and Pointers The Brake System Temperature Scale and Pointers are non-linear and are divided into two segments. The bottom of the scale is the normal region, with values from < 250 C for the 170, and 0 C to < 233 C for the 190. The top of the scale is the amber region, ranging from 250 C to 420 C for the 170 and 233 C to 420 C for the 190. In the amber range the scale is amber and in the normal range the scale is white. The upper 50 % of each scale represent the anber range and the lower 50 % represent the normal range. There is an amber tic mark displayed at 250 C (233 C) Each scale has two pointers, one each to indicate inboard and outboard Brake System Temperatures. In the normal region, the pointers are green. In the amber region, pointers are displayed as solid amber. Brake Temp. data is updated once every 4 seconds and 4 seconds of the same value is displayed before another update.
Figure 6: MFD - System Status Synoptic Page
Crew Oxygen Level Status
Engine Bleed Pressure Digital Readouts
Flight Status
Battery Status
Engine Oil Level Status
Brake Status
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Doors and Access Panels Status The Doors, Access Panels and Door Messages are shown in a plane view image of the aircraft as shown. The upper right section of the Status Synoptic reflects Doors and Access Panels status. The status of each door and access panel can be determined by the appearance of the icon. Doors and Access Panels are displayed in green when closed. If a door or panel is open, it is displayed in red or amber and a corresponding status message is displayed in the upper right corner of the Door and Access Panel window. If the state of a door cannot be determined, the icon is displayed in amber dashes.
Figure 7: MFD - System Status Synoptic Page
Electronic Bay Access Door Fwd Fuselage
Passenger Door Fwd Fuselage
Service Door Fwd Fuselage Fuselage Baggage Door Fwd
Electronic Bay Access Door Central Fuselage
Refueling Bay Access Door Service Door Rear Fuselage Baggage Door Rear Fuselage
Passenger Door Rear Fuselage
last update: Nov06
Hydraulic Bay Access Door
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Airplane Graphic Description: The Flt Ctrl synoptic page contains an Airplane Graphic. The Airplane is a static display that shows the location of some flight control surfaces, status of the Flight Controls actuators and Flight Controls mode of operation.
Hydraulic Synoptic Page The Hydraulics System is comprised of three active and totally independent hydraulic fluid systems. For each of the hydraulic system, a box encloses the fluid quantity readout with its vertical scale / pointer, the system pressure readout with its vertical scale / pointer, and the fluid temperature readout.
MF Spoiler 3 MF Spoiler 5 Status Separation Visual Aid Line
Actuator Status Annunciations Axes Mode Annunciations
Hydraulic System 3 Pressure, Temperature, Quantity Hydraulic System 1 Pressure, Temperature, Quantity
Hydraulic System 2 Pressure, Temperature, Quantity
Hydraulic System Pumps Engine Firewall Shut Off Valves
UR2 User Annunciations UR1 User Annunciations
UL1 User Annunciations
last update: Nov06
Flow Lines UM1 User Annunciations
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UM2 User Annunciations
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Fuel Synoptic Display
The segments are divided as follows for the 190:
The fuel synopticpage contains symbols represent the fuel system component status keeping the basic concept of the system architecture.
Segment
SCALE POSITION
Range - KG
Range - LB
Fuel Tank Quantity Scale and Pointer
Segment 1
40% - 100%
6500 >Tank Qty > 800
14330 > Tank Qty > 1760
The Fuel Tank Quantity Scales are displayed above the readouts. The Fuel Tank Quantity Pointer on each scale is displayed the same color as the segment it is indicating. The scales are divided into three segments. There are tic marks located at the 0%, 20%, 40% and 100% locations on the scales. The segments are divided as follows for the 170:
Segment 2
20 % - 40 %
800 > Tank Qty > 400
1760 > Tank Qty > 880
Segment 3
0 % - 20 %
400 > Tank Qty > 0
880 > Tank Qty > 0
Segment
SCALE POSITION
Range - KG
Range - LB
Segment 1
40% - 100%
4700 >Tank Qty > 600
10361 > Tank Qty > 1320
Segment 2
20 % - 40 %
600 > Tank Qty > 300
1320 > Tank Qty > 600
Segment 3
0 % - 20 %
300 > Tank Qty > 0
660 > Tank Qty > 0
Electrical Synoptic Display There are two Engine Driven Generators displayed on the Electrical synoptic page in the top left and right corners of the display. For each Engine Driven Generator an icon will be displayed indicating wether the generator is or is not functioning. The Electrical synoptic page also displays the Engine Generator Voltage, Frequency, and Load information.
Engine Driven Generator 2 Digital Readout RAT Voltage Display
RAT Icon Transformer Rectifier Unit (TRU) 1 Digital Readout
TRU ESS Digital Readout
Transformer Rectifier Unit (TRU) 1
Transformer Rectifier Unit (TRU) Essential (ESS)
Electrical DC Battery 1 Digital Readout Electrical DC Battery 1
last update: Nov06
Power Unit (APU) Driven Generator
FOR TRAINING ONLY - Reproduction Prohibited
Transformer Rectifier Unit (TRU) 2 Transformer Rectifier Unit (TRU) 2 Digital Readout RAT Frequency Display Electrical DC Battery 2 Electrical DC Battery 2 Digital Readout
DC Ground Power Unit (GPU)
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Environmental Control System Synoptic Display
Anti - Ice Synoptic Display
The ECS synoptic page contains some static objects that help to organize the presentation of various ECS status displays.
The Anti - Ice synoptic page contains some static objects that help to organize the presentation of various anti - ice status displays.The Bleed Air System valves are displayed on the Anti - Ice synoptic page as graphical valve icons. The status of the valves can be determined by the appearance of the icons.
Cockpit / Cabin Temperatures The Cockpit / Cabin Temperatures Display shows the settable and actual temperatures for the cockpit, the forward cabin, and the aft cabin. The Cockpit / cabin Temperatures are located on the ECS synoptic page inside the Aircraft Icon. Centered at the top of the display is the label TEMP C. Below the TEMP C label are the numerical temperature readouts. The set temperatures are arranged vertically on the left, under a label SET. The actual temperatures are arranged vertically on the right, under a label ACTAL. A cutout box surrounds each readout. Centered between the Set and Actual temperature readouts are labels denoting temperature zone: CKPT for cockpit temperature, FWD CAB for forward cabin temperature, and AFT CAB for aft cabin temperature. The Set Temperature readouts provide selected temperature information to the crew. The Actual Temperature readouts provide actual temperature information to the crew.
Bleed Air System Valves The Bleed Air System valves displayed on the Anti - Ice synoptic page include the Engine Inlet Anti - Ice SOVs, HPSOVs, Manifold PRSOVs, Slat Anti - Ice SOVs, a Bleed Isolation Valve, and an APU SOV. Under The Bleed Isolation Valve, the label XBLD is displayed.
Figure 10: MFD - ECS (Environmental Control System page) / Anti - Ice Synoptic Page
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Engine Maintenance Synoptic Display The Engine Maintenance Synoptic page displays dispatch level limitations for each engine. In addition, the engine exceedance information and FADEC channel faults will be displayed. The synoptic page utilizes the entire display. At the top of the Engine Maintenance Synoptic is displayed the title ENGINE MAINTENANCE.
Engine Fault Codes Display Description: The Engine Maintenance page will be displayed a maximum of 14 fault codes per FADEC channel. If more than 14 faults are active, the additional fault codes will be displayed only when actively displayed fault codes are cleared (removed). The fault code data will be displayed in the following format: < FADEC < channel designation >> < fault code> <..>
Dispatch Limitations Display Description: The status of the dispatch limiting faults will be displayed on the Engine Maintenance synoptic page to provide maintenance the engine dispatch limitation status. The left engine dispatch limitations are displayed on the left side of the window while the right engine dispatch limitations are display on the right side.
Engine Exceedances Display Description: When an exceedance is indicated for a specific engine, the Engine Maintenance Synoptic displays the data for seven (7) engine related parameters (N1, N2, ITT, N1 Vibration, N2 Vibration, Oil Temperature and Oil Pressure). The left engine exceedance information is displayed on the left side of the window while the right engine exceedance information is displayed on the right side. For each engine, the exceedance information is displayed in the following format: < exceedance condition > < peak value > < duration >
Figure 11: MFD - Engine Maintenance Synoptic Page
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Maintenance Synoptic Page The Maintenance Synoptic Page is only accessable through the right hand MFD. This page gives the possibility to access Maintenance Messages, System Diagnostic, Tests and other maintenance pages.
Figure 12: MFD - Maintenance Synoptic Page
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The Engine Indicating and Crew Alert System (EICAS) The Engine Indicating and Crew Alert System (EICAS) display presents engine indications, system parameters and Crew Alerting System (CAS) messages. The engine instrument display consists of N1 display, ITT display, N2 display and fuel flow display. Separated but appendant to engine indicating there is the fuel quantity display, the oil display, and the engine vibration display. There is the slat/flap/speed brake display, the trim display, the cabin display, the Auxiliary Power Unit (APU) display, and the landing gear display. The Crew Alerting System (CAS) message window provides the pilots with displayed alerts. Remark: In case of failure in the EICAS display, its information may be presented in the MFD by appropriately setting the reversionary panel.
Figure 13: EICAS
Engine Indicating and Crew Alerting System (EICAS)
Engine Indications System Parameters
Crew Alerting System (CAS) Messages
Remark: In case of failure in the EICAS display, its information may be presented in the MFD by appropriately setting the reversionary panel.
last update: Nov06
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31-61 EICAS Full Panel Introduction FULL PANEL:Pushing the declutter override button displays all of the decluttered items. During an abnormal condition of a system, the affected decluttered item is displayed.
Figure 1: Full Panel
EICAS FULL PANEL
EICAS FULL
Honeywell
DISPLAY
SPLAY
MAX THRUST SET CLB CLOSE
W
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TOW
D E P L O Y
D E P L O Y
FULL
FULL
IDLE
RAT MANUAL DEPLOY
MAX REV
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Declutter (continued) The following items are decluttered during cruise: Oil pressure, Oil temperature, Low pressure vibration, High pressure vibration, Gear position, Flap position, Slat position, Speed brake position
Figure 2: Declutter (continued)
The following items are decluttered during cruise:
Oil pressure Oil temperature Low pressure vibration High pressure vibration Gear position Flap position Slat position Speed brake position
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31-62 Cursor control System Cursor Control Devices (CCD) The Cursor Control Devices (CCD) are instruments installed on both sides of the control pedestal panel. They allow the flight crew to quickly select the displays and position the cursor within the different selectable menus. The CCDs are primarily used on the MFD for selections of formats and functions. They also provide radio tuning on the PFD and EICAS message scrolling on the EICAS display.
Figure 1: Cursor Control Device
Cursor Control Devices (CCD) Cursor Control Devices (CCD) Honeywell
DISPLAY
LOC
APB ATB
125
GS
16O 15O
Pl an
TCAS
WX
Sys t ems
Fuel
4OOO4 R 1O
14O
125
Map
35OO 1OOO
2 1
1O
35OO
AP -3.O
11O 1OO
1O
RA
ILS1 5.3
3OOO4
RA 5OO
5OOM GSPD 3OOKT
1 2
1O
5OO MIN
9O
HDG
33O
36O
29.92
IN
CSR
CHR
36O
O8:12
S
NM
VOR1 VOR2 VHF1 118 O25 119 O25
NAV1 118 O3 119 O3
Checkl i s t
T/O CONFIG Honeywell
STEEP APPROACH
DISPLAY
EICAS DECLUTTER
Honeywell
DISPLAY
MAX THRUST SET CLB CLOSE
W
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TOW
D E P L O Y
D E P L O Y
FULL
FULL
IDLE
RAT MANUAL DEPLOY
MAX REV
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CCD Details The Cursor Control Device (CCD) controls in detail are: The tuning knobs select the value or mode in the data field enclosed by the cursor. Three format locating buttons; the left button places the cursor on the associated PFD, the centre button places the cursor on the associated MFD, the right button places the cursor on the EICAS. The TOUCH PAD is used to direct cursor. The ENTER KEYS are used to select the soft keys.
Figure 2: Cursor Control Device
Cursor Control Device (CCD) Details (Valid for pilot's side only) Tuning knobs Honeywell
31-30 -01 Recorders identification, location and Interface General The Digital Voice Data Recorder (DVDR) system is capable of receiving, recording, and preserving all required flight data parameters and voice recordings on board the aircraft. It combines the function of a Flight Data Recorder with the function of a Cockpit Voice Recorder in a single solid-state recording unit. The Digital Voice Data Recorder Unit is capable of recording the most recent: 120 minutes (2 hours) of Audio information from 4 analog input channels; 25 hours of Flight Data information received via an ARINC 717 Bus; 120 minutes (2 hours) of Digital Communication Messages, and Greenwich Mean Time (GMT) from an ARINC 429 source.
Figure 1: DVDR
The Digital Voice Data Recorder (DVDR) System iving rding serving DVDR
Cockpit Voice Recorder
Flight Data Recorder
-120 min (2 hours) of audio information
4x analog input channels ARINC 717 Bus
-25 hours of Flight Data information
ARINC 429 Bus
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DVDR Components The Digital Voice Data Recorder System has the following components: • Two identical Digital Voice Data Recorder units; DVDR 1 is installed in the forward E Bay, DVDR 2 is installed in the aft E Bay. • Two Impact Switches, one each mounted next to its respective Digital Voice Data Recorder. • One Tri-axial Accelerometer, installed in the left hand side wingstub area. • One DVDR Control Panel, located in the cockpit on the overhead panel. • One Cockpit Area Microphone, mounted on the frame between both windshields. • Ten Load Cells (force sensors), four of which are mounted directly in the interconnection rods for the Rudder Pedal Systems, two in the interconnection rods for the Elevator Control Columns, and four installed in the command cables for the Aileron Control Wheels.
Figure 2: DVDR Components ELEVATOR
LOAD CELL
C
C
AILERON
C
DVDR 1
LOAD CELL
C
C
RUDDER - LOAD CELL
C
Impact Switch
C LOAD CELL
C
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The DVDR Audio Section The audio section of the Digital Voice Data Recorder has two parts. The primary crew audio input signals are transmitted by three narrow band voice channels, delivered from the three Audio Panels. The Cockpit Area Microphone (CAM) picks up the acoustical environment of the flight deck, and through a pre-amplifier in the Cockpit Control Panel, the signal is delivered using a single wide band channel to the recorder. Thus, there is a total of 4 analog Audio channels. Audio Monitor signals from the DVDRs are sent to the DVDR Control Panel. A standard phone jack on the Cockpit Control Panel allows audio monitoring using headphones.
Figure 3: The DVDR Audio Section
TEST
DVDR CONTROL PANEL D
FDR
F FT
STATUS
DVDR TEST
CVR ERASE HEADPHONE
Overhead panel
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31-30-02 Recorders Component Description and Operation Introduction Each Digital Voice Data Recorder includes three Shop Replaceable Units (SRU) and an Underwater Locating Beacon (ULB). • First there is an Interface and Control Board (ICB), with PCMCIA Interface for Flight Data Down load. • Second there is a Crash Survivable Memory Unit (CSMU). The CSMU is a solid state, non-volatile, mass storage device enclosed in a protective case. It provides storage for all input data and also ancillary system data. • Third there is a Power Supply (PS) Assembly, which converts +28 VDC aircraft power to secondary power for the SRUs and provides power-on reset, power failure monitoring, and significant power hold-up capability. The DVDR unit is cooled by convection and radiation to ambient air. No forced-air cooling is required. The Underwater Locator Beacon (ULB), is an aid to underwater retrieval. The ULB is mounted directly in front of the CSMU to minimize the likelihood of detachment in case of an incident. This mounting position allows the ULB to double as a handle for the unit, and provides easy access to the battery. The battery replacement date is written on a label on the outside of the ULB. The lifetime of the internal battery is six years.
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Page 1
Figure 1: Digital Voice Data Recorder components
Underwater Locating Beacon (ULB) Crash Survivable Memory Unit (CSMU) - solid state - non-volatile - mass storage - storage for all input data - storage for ancillary system data
- Power Supply (PS) Assembly power-on reset power failure monitoring power hold-up capability
Interface and Control Board (ICB)
The life time of the internal battery is six years.
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The Digital Voice Data Recorder Control Panel The Digital Voice Data Recorder Control Panel uses standard input/output interfaces for a Cockpit Voice Recorder and a Flight Data Recorder. The Recorder Test button (DVDR TEST) is used to initiate a Cockpit Voice Recorder and Flight Data Recorder test within both DVDRs simultaneously. You can switch the FWD/AFT Switch position to check the test status for either DVDR 1 – switched to FWD, or DVDR 2 – switched to AFT. The CVR ERASE button initiates the Cockpit Voice Recorder Erase Function. A HEADPHONE jack is available for audio monitoring. The position of the FWD/AFT Switch also determines the CVR audio monitor source - forward or aft DVDR. The Cockpit Area Microphone (CAM) picks up area audio and delivers it to the preamplifier of the DVDR Control Panel.
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Figure 2: The Digital Voice Data Recorder Control Panel
COCKPIT AREA MICROPHONE
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System architecture Flight data is received from the Modular Avionics Unit (MAU) via an ARINC 717 interface: the Custom input/output Module. The Custom I/O Module accepts data via the Avionics Standard Communications Bus (ASCB) from a multitude of sources, and converts it to the ARINC 717 format. Important: Both ARINC 717 buses send the same data to both recorders. A source selection logic software within the Custom I/O Module determines what data to pull from the ASCB. The function is common to both the DVDR 1 and DVDR 2. However, the Custom I/O Module for Recorder 1 is in Modular Avionics Unit 1, and the Custom I/O Module for Recorder 2 is in Modular Avionics Unit 3. Other data transmitted via the Custom I/O Module includes: A discrete "Flight Data Recorder FDR enable" to the Recorder; The Greenwich Mean Time (GMT) via an ARINC?429 Bus to the Recorder, and Central Maintenance Computer Data via an ARINC?429 Bus to and from the Recorder.
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Figure 3: System Architecture
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System Architecture (continued) A Cockpit Voice Recorder Erase, discrete from MAU 1, is supplied to the DVDR Control Panel. The CVR erase switch on the DVDR Control Panel is used for manual CVR erasure. Note: Analog audio is erased simultaneously on both Digital Voice Data Recorders. The Flight Data Recorder and the Cockpit Voice Recorder sections can be tested at the same time, as well as both DVDRs. The Processor Modules in MAU 1 and MAU 2 host the Monitor Warning Function (MWF). The MWF monitors the status of the DVDR System and puts relevant Crew Alerting System (CAS) messages on the Avionics Standard Communications Bus. CAS messages will be displayed on the Engine Indication and Crew Alerting System Display. The Communication Management Function (CMF) is hosted in the Processor Module of the MAU. The CMF pulls Digital Audio Communication Messages from the ASCB and makes them available for the DVDR via an ARINC 429 bus. For DVDR System 1 the Processor Module of MAU 1 sends the Digital Audio Communication Messages to DVDR 1, and for System 2 the MAU 3 sends the messages to the DVDR 2 via the ARINC 429 bus.
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Figure 4: System Architecture (continued) MAU 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B
16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9
8 7 6 5 4 3 2 1
2 B
C H
B U
#
S
CMC GPS 1 Power Supply 2 ESS 1 FCM 1 A 1
16 15 14 13 12 11 10 9
AIOPB1 PROC 1 NIC 1 (A) (ID = 1) FCM 2
AIOPA1
A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1
2 B
SPARE SPARE GENERIC I/O 2
U
C H
8 7 6 5 4 3 2 1
2 B 2 B
B
#
MAU 3 C H
B U
U S
C H
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DATABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2 Power Supply 1 DC 2
#
S
A 1 2 B 2 B
A 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1
2 B 2 B 2 B
Power Supply 2 ESS 2/DC 2 BRAKES (INBD) CONTROL I/O 2 AIOPA2
B S
CUSTOM I/O 1
CONTROL I/O 1 BRAKES (OUTBD) PSEM 1
B U S
C H
16 1 B 15 14 13 12 1 B 11 10 1 B 9 1 B 1 B
A 1 A 1
A 1 A 1 A 1 A C H
8 7 6 5 4 3 2 1
1 B
1 B B
U
U
#
U S
C H
Power Supply 1 ESS 1
C H
B U S
C H
B U S
A 2 2 A
FCM 3 A 2 GENERIC I/O 3 A 2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 NIC 5 (A) (ID = 33) CUSTOM I/O 2
MWF - Monitor Warning Function Tri - Axial Accelerometer - Pilots load cells - Copilot load cells Custom I/O 1 - Data to DVDR 1 last update: Dec06
Custom I/O 2 - Data to DVDR 2
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The Tri-axial Accelerometer The Tri-axial Accelerometer is a hermetically sealed instrument that simultaneously measures acceleration along three axes: • longitudinal, • lateral and • vertical. The accelerometer consists of three separate seismic sensors that respond to the force along these three axes. The Impact Switches removes electrical power to both mounted DVDR units when the aircraft is impacted. Load Cells are force sensors that collect flight data cockpit control forces. They measure pilots' and copilots' forces on the airplane control stick, wheel and pedals, and provide required flight data parameters via the MAUs to the DVDRs.
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Figure 5: The Tri- Axial Accelerometer ZONES 143 144
SPAR II SPAR III
A B
A
TRIAXIAL ACCELEROMETER
RIB 1 (REF.)
C
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Start Logic The DVDR will automatically start recording CVR data as soon as power is applied to the unit, and will continue until power is removed from the unit. The unit will automatically begin recording FDR information when the engines are turned on, or the aircraft is in the air (WOnW false). The FDR can also start recording for maintenance purposes by selecting the FDR recording selection on the MCDU. This logic is depicted in the following start logic diagram. To perform a test of the DVDR operation, the pilot can press and hold the DVDR Test button on the Control Panel. This test may be performed both on the ground and in flight. The system will perform a self-test and respond with an aural indication if the test is successful.
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Figure 6: Start logic
MCDU recording selection
FDR Start/Stop
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MCDU The pilot input to the DVDR is done through the DVDR control panel and MCDU (Multifunction Control Display Unit). In the MCDU MISC MENU page, a DVDR menu selection is supplied. This menu selection is available only when WOW is true, both engines are powered off (engine running not indicated) and both engines starter valve is closed. If any of these conditions are false, the DVDR is assumed to be recording and the DVDR selection is not shown on the MISC MENU page. The DVDR page contains the menu selection that starts the FDR recording function (used for maintenance purposes). The LSK (Line Select Key) 1L is pushed to select the FDR recording to ON or OFF. This selection affects both the DVDRs identically. The record function is on or off for both the FDRs.
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Figure 7: Digital Voice - Data Recorder (DVDR) System
DVDR LSK 1L
F D R
ON
1 / 1
R E COR D I NG O F F
R E T U R N
A PERF
NAV
PREV
FPL
PROG
DIR
BRT
CB
DIM MENU
DLK
NEXT
TRS
RADIO
A
B
C
D
E
F
G
H
I
J
K
L
1
2
3
+/ -
M
N
O
P
Q
R
4
5
6
/
S
T
U
V
W
7
8
9
X
Y
Z
DEL
CLR
SP
0
MCDU
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Hand Held Download Unit FDR data may be downloaded while still installed in the aircraft via the download connector interface using the Hand-Held Download Unit (HHDLU) or other suitable FDR download equipment. The HHDLU also provides a DSDU (Data Signal Display Unit: real time data display) function where the FDR data may be displayed concurrently as it is being recorded. The FDR Download (and DSDU) function is activated when the GBE Present discrete from the 28-pin download connector is grounded. This interface also provides power to the HHDLU. When the HHDLU is hooked up to the DVDR, the FDR data can be downloaded with the FDR not enabled. The HHDLU is connected to the download connector of the DVDR, allowing the unit to remain installed in the aircraft and connected to aircraft data sources via the main aircraft connector. In DOWNLOAD mode, the flight data from the DVDR unit is copied from the crash survivable memory module onto the cartridge memory installed in the HHDLU. Once the DOWNLOAD is complete, the cartridge can be removed from the HHDLU and installed into a ground-based personal computer for further processing and data distribution.
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Figure 8: Hand Held Down Unit
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PCMCIA Interface To allow rapid flight data download with minimal operator interface, the integral PCMCIA Interface will allow a rapid data transfer of the complete flight data memory contents to an extractable ATA-Type PCMCIA card. This will allow the complete 25 hours of encrypted flight data contents of a 256 wps (36 megabytes) to be transferred quickly. The PCMCIA Interface provides a Memory Cartridge slot that accepts removable PC Cards for the down loading of the DVDR unit encrypted Flight Data. Cartridge access is provided with a simple door cover that prevents debris from inadvertently entering the cartridge slot. The cartridge insertion and removal is assisted by an ejection mechanism, which provides a simple pushto-release fingertip actuation. The cartridge is ejected a sufficient distance to allow the operator to grasp when released.
• Once the flight data transfer from the crash survivable memory to the PCMCIA card begins, the BUSY (Transfer in Progress) LED will illuminate and all recording operations (both cockpit voice and flight data) will be suspended. The BUSY LED will remain illuminated until all of the flight data has been transferred to the PCMCIA card. If the BUSY LED does not illuminate, then either the recorder does not recognize that the PCMCIA card has been inserted, or power has not been applied to the recorder. Once the transfer process is completed, the BUSY LED will extinguish and the DONE (Transfer Complete) LED will illuminate for a maximum time of 30 seconds, or until the PCMCIA card is ejected. The PCMCIA card is ejected by depressing the EJECT button located immediately to the left of the PCMCIA slot.
The process to download flight data via the ATA-Type PCMCIA card is as follows: • With the recorder powered “ON”, the PCMCIA card is inserted into the PCMCIA slot, which is accessed from the front panel via a protective cover. • The PCMCIA card is verified to contain sufficient available memory to perform the transfer of the entire flight data contents of the DVDR unit. If not enough physical memory exists on the card (i.e capacity is less than the DVDR unit memory content or the card cannot be accessed, a failure indication will be noted (simultaneous flashing of both the BUSY (Transfer in Progress) and DONE (Transfer Complete) LED’s on the DVDR front panel). • If sufficient free memory space is available, a new file is automatically created.
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Figure 9: PMCIA Interface
1. EJECT BUTTON
2. PCMCIA CARD SLOT
BUSY DONE
3. DVDR UNIT BUILT-IN TEST EQUIPMENT INDICATOR
BITE
GSE DOWNLOAD J2
6. RS-422 DOWNLOAD CONNECTOR
5. TRANSFER IN PROGRESS (BUSY) INDICATOR
4. DOWNLOAD COMPLETE INDICATOR
DVDR UNIT FRONT PANEL
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31-32 Quick Access Recorder System Introduction The QAR (Quick Access Recorder) is a piece of equipment that records a large amount of data which can be ARINC (Aeronautical Radio Incorporated) - 717 data or ARINC-429 data. It allows easy access to recorded data by quick removal of the QAR card. The QAR is installed on the equipment rack in the aft avionics compartment.
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Page 1
Figure 1: Quick Access Recorder System Overview
A717 DATA
DVDR
MAU 3
A429 DATA
QAR
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28VDC
DC BUS 2
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General Description The QAR has two main blocks: • The AFDAZU (Auxiliary Flight Data Acquisition Unit), responsible for determining the incoming transmission rate and for collecting the data from the ARINC-717 and making it available for the DMU (Data Management Unit) part. • The DMU, responsible for collecting the ARINC-429 data, composing it with the ARINC-717 data and recording it into the QAR card. If one of the two blocks fails, it does not implicate that the other will fail too. If the ARINC-717 acquisition function of the AFDAU fails, ARINC-717 data will be lost, but the QAR record ARINC-429 data normally. If the DMU fails in acquiring ARINC-429 data, the QAR will still record data from ARINC-717.
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Figure 2: QAR-Block Diagram
A717 DATA
DVDR
MAU 3
A429 DATA
GENERIC I/O 3
A429 DATA
CUSTOM I/O 2
DC BUS 2
AFDAU
28VDC
DMU PCMCIA
QAR
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Components QAR UNIT The QAR unit consists of a LRU (Line Replaceable Unit) comprising a chassis closed by two side panels. Openings on the top and lower panels ensure air cooling capability. A connector attached to the back panel allows connection with the aircraft systems. The main components of the QAR unit are: • An access door to a connector, used for QAR card access. • One identification label. • Three indicator status red lights (DMU FAIL, FDRS FAIL and AFDAU FAIL). • Two status LEDs (a green color and an amber color). QAR CARD The QAR card is a standard PCMCIA (Personal Computer Memory Card International Association) card (PC (Personal Computer) Card) capable of holding more than 70 hours of aircraft operation.
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Figure 3: QAR - Component Location
QAR UNIT
A ACCESS DOOR
IDENTIFICATION LABEL
C
QAR CARD SLOT
INDICATION STATUS LIGHTS
B
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QAR CARD STATUS LEDS
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Operation
ory space dedicated for each one is programmable by GSE. Two recording modes can be selected by GSE:
The QAR is energized as soon as the Aircraft is powered ON. The unit starts and stops recording in accordance with the START/ STOP logic programmed by the operator through the software tool named Ground Support Equipment (GSE). There are three red lights on the unit front panel indicating the system status. No provision has been made to warn the system status to the crew.
• Continuous Mode: data is continuously recorded on the QAR card. When it is full (i.e. the allocated memory space is full), the oldest data is deleted, allowing the recording to continue; • Simple Mode: data is continously recorded on the QAR card. When then it is full (i.e. the allocated memory space is full), the recording stops.
Once energized, the AFDAU will recognize the aircraft type through the aircraft ident descrete pins, and then will select the data frame in accordance with the aircraft type. If the AFDAU does not recognize the aircraft installation, the AFDAU Fail light is turned ON. The QAR unit global status is displayed by the following three LEDs located on the front panel board:
The QAR is fed with +28 VDC (Volt Direct Current) from DC (Direct Current) bus 2 and is equipped with power voltages monitoring, transients and restart circuitry, maintenance, two +5 VDC segregated outputs (potentiometers) and three discrete outputs: AFDAU Bite Out, FDR System Bite Out and DMU Bite Out.
• FDRS FAIL: lighted when no ARINC-717 input is detected and when the AFDAU fails. • AFDAU FAIL: lighted when the AFDAU fails. It indicates no QAR installed or inappropriate installation, failure in detecting the transmission rate or internal failures. • DMU FAIL: lighted when DMU part of the QAR fails. The fails can be in the ARINC-429 data acquisition, battery failures or internal failures. The power supply unit delivers the internal voltages required for the functioning of the equipment. The QAR will start- and stop- recording in accordance with the logic programmed by the operator at the GSE. The default logic delivered is: Left N1 (Fan Rotor Speed) > 10% or Right N1 > 10% or Left N2 (Core Rotor Speed) > 10% or Right N2 > 10% or Air/ Ground = FALSE. Basically, two kinds of data frames can be recorded at the QAR card: an exact copy of the FDR (Flight Data Recorder) data frame, at 256 WPS (Words per Second) and a programmable data frame, with parameters and rate programmable by operator. The rate of the programmable frame can be selected from 64 WPS to 2048 WPS. Both data frames can be recorded in the QAR card. The mem-
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Figure 4: QAR - Controls/ INDICATORS
A
QAR UNIT FRONT PANEL
C
AFDAMU (AFDAU+DMU) P/N
F6151 75512 PARIS CEDEX 15 FRANCE Made in France
ED35E109-04-00
Amdt S/N 483
INSP
DMU FAIL
Date 01/04 Weight
4,2Kg
FDRS FAIL AFDAU FAIL
INDICATION STATUS RED LIGTHS
C
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Training Information Points STATUS LEDS INFORMATION When you insert the QAR card to the QAR card slot, the green color status led blinks at a frequency of 1 Hz approximately, as a manner of recognize the card. If the QAR card is empty, the green color status led will go out. If there is a database in the QAR card, this database will be loaded to the QAR unit and the green color status LED will blink at a higher frequency during the loading procedure. The green color led stays on after the loading procedure until the QAR is powered off. If the QAR card is not correctly formatted, the amber color status led comes on.
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31-41 MAU Modular Avionics Unit Introduction There are three MAU’s installed in the Embraer 170.The three MAU’s contain a large number of avionics systems. Using various communication buses the equipment inside the MAU’s can communicate with all the A/C systems.
Figure 1: Primus Epic MAU Schematic
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Modular Avionics Unit (MAU) Each MAU is a metal cabinet solidly grounded to the aircraft frame. Each cabinet contains slots that line replaceable modules, (LRMs) are plugged into to make contact with a virtual back plane bus for power and communication. Shielded connectors on the front of the modules provide protection against the effects of lightning strike or HIRF. As there are typically two or more modules for each function, the power source for each module is chosen to be different from the other to further ensure function availability. Cooling fans on the rear of the MAU eliminate the need for forced air cooling. There are two 16- slot MAUs and one 20- slot MAU in the ERJ 170.
Module Functions The functions of the modules within the ERJ 170 MAU are as follows although more detailed descriptions may be found in the respective SDR for a particular function. These are descriptions only for the types of modules that are used as generic resources to other system components that perform aircraft functions.
Figure 2: MAU
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General Description The ASCB supplies an aircraft-wide communications path between the MAU, PFD, MFD, EICAS, SPDA, and MRC. Each unit uses a common interface, called the NIC (Network Interface Controller), to send and receive data over the ASCB. In the case of the MRC, the NIC function is done by a NIM (Network Interface Module). The NIC acts as a gateway for modules to access the ASCB. Basically, the NIC supplies the interface between the backplane bus on each unit and external ASCB. The NIC has functions that keep the traffic on the data bus synchronized. The NIC also controls data transmission to and from the LAN (Local Area Network). The LAN is an Ethernet-based LAN used for data loads, maintenance, and test purposes. The LAN connects to the same units as the ASCB. Each MAU has an in-line bus coupler connected to a NIC between the MAU and the cross-side ASCB. The bus coupler isolates the cross-side ASCB from an MAU failure.
Figure 3: Modular Architecture
General Modular Avionics BIC BIC
- power supplies
BIC
- backplane (motherboard)
USER M odule USER M odule
BIC
Two channels per cabinet
NIC M odule
USER M odule
BIC
Modular Avionics Unit (MAU):
USER
BIC
PWR M odule
NIC M odule
- one Network Interface Controller - various user modules
Each NIC is connected to: - on- side primary bus - on-side backup bus - cross- side primary bus
A S C B
L A N
M odule
PWR M odule
Dua l C ha nnel MAU
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ASCB Bus System Introduction The ASCB (Avionics Standard-Communication Bus) is the system data bus that moves data among the avionics and (secondary) power distribution assemblies. The ASCB supplies the wiring and connections for two-way data transfer between the units at 10 Mb/s rates. The ASCB connects the MAU (Modular Avionics Unit)s in the forward and middle avionics compartments with these other major units: • • • • •
PFD (Primary Flight Display)s MFD (Multi-Function Display)s EICAS (Engine Indicating and Crew Alerting System) MRC (Modular Radio Cabinet)s SPDA (Secondary Power Distribution Assembly)s
The ASCB is made up of two primary and two backup buses. Each unit with the ASCB service connects to both the left and right primary buses. Each unit also connects to either its left or right back-up buses. This arrangement supplies both redundancy and fault isolation. The physical and electrical separation of the backup buses decreases the likelihood that one defective LRU (Line Replaceable Unit) will disrupt communications on the bus.
Figure 4: System Network Buses
ASCB - D
4 BUSES FOR REDUNDANCY REQUIREMENTS LH (Backup) LH (Primary)
PFD
MFD
EICAS
MFD
MAU 1
Channel A Channel B
PFD
MAU 2
RH (Primary) (P i )P i RH (Backup)
Channel A Channel B
LAN
MAU 3
Channel B Channel A
MRC 1
A (Left)
MRC 2
A (Backup) B (Right) SPDA 1
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SPDA 2
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31-51 Aural Warning Function General Aural warnings are used when the pilots require immediate knowledge of the condition without having to look at a visual display or indicator. Aural warnings are alert tones, horns, bells, clicks, beeps and voice messages. The electronic display system has two aural warning drivers, which are responsible for generating and prioritizing aural warnings. Aural warnings sound in a sequence, and are never truncated. Aural warnings are cancelled automatically when the alerting situation no longer exists, or when they are reset manually by the pilot. In the event of multiple alerts, the highest priority alerts sound first.
General Description TABLE - AURAL WARNING FUNCTION - AURAL WARNINGS
The Aural Warning System has two aural warning drivers and two MWF (Monitor Warning Function)s. The aural warning drivers supply aural messages. The MWFs supply aural warning logic control. The MWF controls all avionics aurals except TCAS (Traffic Alert and Collision Avoidance System). The MWFs are present in the two processor modules on different MAU (Modular Avionics Unit)s (MAU1 and MAU2). The aural warning hardware is on the control I/O (Input/Output) module. All MWF aurals are based on a priority. The MWF sequences through the active MFW aurals, and starts with the aurals that have the highest priority. If a higher priority MWF aural becomes active while a lower priority aural is being played, the higher priority aural is played after the lower priority aural ends. Each aural warning is aurally distinct from all the other warnings. The voices are clear and they use full words (they do not use abbreviations that are used with any related visual message). There is a silent interval (pause) between different aural warnings. When only one aural warning is active, a silent period follows the repeated single warning to make sure the repeated audio warning does not distract the pilots. The aural warnings are heard in a monotone female voice.
TABLE - AURAL WARNING FUNCTION - AURAL WARNINGS (Continued)
PRIORITY
ALERT
TONE/ VOICE MESSAGE
TYPE
CANCELABLE
1
Autothrottle (normal or abnormal)
²Throttle²
Continuous
Yes
²Autobrake²
Single
No
²Takeoff OK²
Single
No
0 Continuous
Continuous
No
No
0
Autobrake Takeoff Configuration
0
Aural Warning A Test
²Aural Warning Test A²
Single
No
0
Aural Warning B Test
²Aural Warning Test B²
Single
No
²ATC Messsage²
Single
No
²T ² (during 7 seconds)
Continuous
No
0
0
CMF (Communications Management Function) Trim Malfunction
[1] EGPWS and TCAS voice messages are not detailed in this section.
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Operation
3 - Emergency
The aurals are used to get the flight crew’s attention. The aurals can be tones or voices. Each aural is necessary for the flight crew to do the correct procedure. If a tone alert and a voice message occur at the same time, the volume of the tone decreases to allow the flight crew to understand the voice message and tone. The aural warning function automatically selects the aurals in order of importance. The aurals are heard in a sequence and are not truncated. The alerts are cancelled automatically when the alert condition no longer exists. If multiple alerts are requested, the alerts with the highest importance (emergency) are heard first. Once all emergency aural alerts have been cancelled or satisfied, then the abnormal alerts are annunciated. If all emergency and abnormal alerts are annunciated. If all emergency and abnormal alerts have been cancelled or satisfied, then the advisory alerts are annunciated. If there are no other alerts, the information alerts are annunciated.
The emergency alerts tell the flight crew of an emergency condition, such as a dangerous aircraft configuration, or a serious system failure. The master warning is repeated with three-second intervals between alerts until the master warning reset switch is pushed. 2 - Abnormal The abnormal alerts are used in conditions (such as system malfunctions) that have no immediate impact on safety. When an abnormal fault occurs, the master caution tone is heard at five-second intervals until the master caution reset switch is pushed. 1 - Advisory
A special condition occurs when a windshear, TCAS, or EGPWS alert occurs. In that condition, no voice messages are supplied to prevent messages from these systems not being heard clearly.
The advisory alerts relate to a condition such as a system malfunction leading to a loss of system redundancy, or an aircraft system not being fully operational. A single alert signal is heard and cancels automatically.
Any alert is heard completely before another alert is heard (even if the priority is higher). When two or more time-critical alerts occur at the same time, or one after the other, they are presented in chronological order. The highest number has the highest importance.
0 - Information
TABLE - AURAL WARNING FUNCTION - Aural Priorities PRIORITY
CONDITION
3
Emergency
2
Abnormal
1
Advisory
0
Information
The information alerts correspond to an information condition. A single alert signal is heard and cancels automatically.
Figure 3: Emergency priority
MAU 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B
16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9
8 7 6 5 4 3 2 1
2 B
C H
B U
#
S
CMC GPS 1 Power Supply 2 ESS 1 FCM 1 A 1
16 15 14 13 12 11 10 9
AIOPB1 PROC 1 NIC 1 (A) (ID = 1) FCM 2
AIOPA1
A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1
2 B
SPARE SPARE GENERIC I/O 2
U
C H
8 7 6 5 4 3 2 1
2 B 2 B
B
#
MAU 3 C H
B U
U S
C H
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DATABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2 Power Supply 1 DC 2
#
S
A 1 2 B 2 B
A 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1
2 B 2 B 2 B
Power Supply 2 ESS 2/DC 2 BRAKES (INBD) CONTROL I/O 2 AIOPA2
B S
CUSTOM I/O 1
CONTROL I/O 1 BRAKES (OUTBD) PSEM 1
B U S
U S
C H
Power Supply 1 ESS 1
C H
B U S
C H
16 1 B 15 14 13 12 1 B 11 10 1 B 9 1 B 1 B
A 1 A 1
A 1 A 1 A 1 A C H
8 7 6 5 4 3 2 1
1 B
1 B B
U
U
#
C H
B U S
A 2 2 A
FCM 3 A 2 GENERIC I/O 3 A 2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 NIC 5 (A) (ID = 33) CUSTOM I/O 2
AURAL WARNINGS DRIVERS - IN CONTROL I/O MODULES AURAL WARNING FUNCTION - IN MAU PROCESSORS
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Intentionally left blank
Figure 4: Aural Warning Function - Block Diagram
AURAL WARNING CONTROL (AWC)
AURAL WARNING CONTROL (AWC)
(PRIORITY)
(NOT PRIORITY)
ASCB - PRIMARY
ASCB - BACKUP
AURAL WARNING HARDWARE ON CONTROL I/O MODULE
AURAL WARNING MUTE
TCAS
AURAL WARNING MUTE
AURAL WARNING HARDWARE ON CONTROL I/O MODULE (SLAVE)
(MASTER) AUDIO AUDIO
AUDIO
AUDIO
MODULAR RADIO CABINET 1
AUDIO
MODULAR RADIO CABINET 2
AIRBORNE AUDIO SYSTEM
COCKPIT SPEAKERS
AURAL WARNING FUNCTION - BLOCK DIAGRAM
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31-51-01 Take-Off Config Warning Takeoff Configuration Monitor The Takeoff Configuration Monitor is a function used on ground to check if the aircraft is configured for takeoff. It is designed to prevent an attempted takeoff of the aircraft without the aircraft properly configured for takeoff. It is automatically activated during the initial portion of the takeoff roll as soon as the flight crew selects takeoff power. The Takeoff configuration monitor checks the Emergency Parking Brake, Spoiler, Pitch Trim and Flap/Slat position and provides visual and aural warnings to the crew as a result of an improper configuration. It can be manually activated via the T/O Config PBA installed in the pedestal. The Takeoff Configuration Monitor indicates true if the airplane is not configured for takeoff, and indicates false if the airplane is properly configured for takeoff. Whenever the aircraft is configured for takeoff and the Takeoff Configuration Monitor is active, no indication is produced. Whenever the aircraft is not configured for take off and the Takeoff Configuration Monitor is active, the monitor is tripped and a warning CAS Message is displayed, an aural warning is sounded and the corresponding display symbology becomes red inverse video. The crew can manually activate a Takeoff Configuration check by pressing the Takeoff Configuration push button in the cockpit. When the check is completed, an aural indication is provided if the aircraft is in takeoff configuration, and a CAS message is provided if the aircraft is out of takeoff configuration.
Emergency/Park Brake Takeoff Configuration The emergency/park brake setting during takeoff factors into the Takeoff Configuration Monitor. If pressure is applied to the Emergency/Park Brake, the aircraft is out of takeoff configuration and the Takeoff Configuration Monitor is tripped.
Spoilers Takeoff Configuration The spoilers position during takeoff factors into the Takeoff Configuration Monitor. All surface panels must be retracted for takeoff and the speed brake lever in the cockpit must be in the stowed position. If the spoilers are not stowed, or are not commanded to stowed by the speed brake lever, the aircraft is out of takeoff configuration and the Takeoff Configuration Monitor is tripped. If any spoiler panel inadvertently opens without crew command (Control Wheel) then the Takeoff Configuration Monitor will trip.
Pitch Trim Takeoff Configuration The pitch trim setting during takeoff factors into the Takeoff Configuration Monitor. The pitch trim must be in the green band on the EICAS pitch trim scale for takeoff. If the pitch trim is outside of the green band, the aircraft is out of takeoff configuration and the Takeoff Configuration Monitor is tripped.
Slat/Flap Takeoff Configuration The slat/flap setting during takeoff factors into the Takeoff Configuration Monitor. For the final aircraft configuration, the flaps/slats must be in position 1,2,3, or 4 for takeoff. If the flaps/slats are in position 0.5 FULL, or in motion (green dashes), the aircraft is out of takeoff configuration and the Takeoff Configuration Monitor is tripped.
Take-Off configuration Check Button A test button is provided to allow checking the takeoff configuration warning integrity, by simulating power levers in the advanced position. A voice message is generated after successful tests. Unsuccessful tests will generate a “NO TAKEOFF CONFIG” message on the EICAS display and a voice message associated with the item out of configuration.
Figure 2: Take-Off Configuration Check Button
EMERGENCY - PRIORITY 3
Order of priority: 1 - Stall 2 - EGPWS 3 - TCAS 4 - Fire
WARN 3 sec interval last update: Jun06
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31-52 Central Warning System (CWS)
(warning, caution, advisory, information/status). When new warning, caution, and advisory messages are received, their status is unacknowledged. The warning and the caution messages continue to change from the inverse video to the regular video until the flight crew reads the alert. The advisory messages automatically change from the inverse video to the regular video after five seconds. The information and the status messages show as acknowledged when initially received, and do not change from the regular video to the inverse video.
Operation CREW ALERTING SYSTEM The CAS display is on the top right corner of the EICAS DU. The CAS display has a message field that shows the new messages, the out-of-view messages, and the CCD (Cursor Control Device) focus. When the CCD focus is set to EICAS, the bottom left part of the CAS window becomes cyan. The CAS shows four types of messages as follows: • The warning messages show in red. These messages show an emergency condition that demands immediate action by the flight crew. • The caution messages show in amber. These messages indicate that the aircraft operation or the condition of an aircraft system is not correct. The flight crew must take immediate action. Other procedures may be necessary. • The advisory messages show in cyan. They indicate aircraft systemsthat need to be monitored by the flight crew. • The information/status messages show in white. They supply acockpit indication on an aircraft system condition, but are not partof the warning system.
OUT-OF-VIEW MESSAGES The out-of-view message field shows the number of messages in each level. The messages show in a specific color. An arrow shows if the message is above or below the window. The out-of-view message display (digits and arrows) flashes continuously when messages that have not been read are outof-view. If there are no messages out-ofview, the message display is blank, except for the END message. The END message lets the flight crew know that there are no more messages. The messages are out-of-view either because the number of active messages exceeds the number of lines in the message field, or because the flight crew has deliberately moved the message out-of-view. The out-of-view message display on the right shows the number of messages that are out-of-view below the message field. The out-ofview message display on the left shows the number of messages that are out-of-view above the message field. The out-of-view message fields can show up to 99 messages. If there are more than 99 related messages, the out-of-view message field remains at 99.
CAS MESSAGE FIELD The CAS message field can show up to 15 lines of text, with a maximum of 22 characters per line. If the message has a related checklist, the left character is an asterisk. All of the CAS messages are justified on the left and start in the second character space (from left to right). The END message is centered. The warning messages show on the top of the message field, followed by the caution messages, the advisory messages, and the information/status messages. The message lines that are not used are shown as blank spaces. The alert messages show from top to bottom in chronological order for each category. A new message shows as the first message of the group
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Chapter 31-52
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Figure 1: Central Warning System (CWS)
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MASTER WARNING/CAUTION INDICATION The two MWFs control the master warning/caution indication. The MWFs are in the two processor modules and operate as priority and non-priority. One of them is located in the MAU (Modular Avionics Unit) 1 and the other is in MAU 2. The MWF that is powered-up first isthe priority MWF. The non-priority MWF operates as a backup. The MWF uses discrete outputs to operate the master warning and the master caution indicators. The light flashes continuously (0.5 second ON and 0.5 second OFF). The master warning or caution shows again with three-second intervals between alerts. The light goes off when the condition stops or when the pilot pushes the master warning/master caution pushbutton.
EICAS MESSAGES INHIBITION CODE
AFTER
BEFORE
DESCRIPTION
K1
Electrical Power ON
1st Engine Started
A/C parked
K2a
1st Engine Started
TLA> TO Power
A/C taxiing
K2b
TLA>TO Power
80 knots
TO Roll
K3
80 knots
400 ft (takeoff)
Takeoff
K4
400 ft (takeoff)
200 ft (landing)
Climb, cruise, approach
K5
200 ft (landing)
5 seconds after touchdown
Landing
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Page 3
Figure 2: Monitor Warning Function
MAU 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B
16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9
8 7 6 5 4 3 2 1
2 B
C H
B U
#
S
CMC GPS 1 Power Supply 2 ESS 1 FCM 1 A 1
16 15 14 13 12 11 10 9
AIOPB1 PROC 1 NIC 1 (A) (ID = 1) FCM 2
AIOPA1
A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1
2 B
SPARE SPARE GENERIC I/O 2
U
C H
8 7 6 5 4 3 2 1
2 B 2 B
B
#
MAU 3 C H
B U
U S
C H
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DATABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2 Power Supply 1 DC 2
S
C H
Power Supply 1 ESS 1
C H
B U S
B U S
C H
16 1 B 15 14 13 12 1 B 11 10 1 B 9 1 B 1 B
A 1 A 1
A 1 A 1 A 1 A C H
8 7 6 5 4 3 2 1
1 B
1 B U
PROC 1 = ADA 1, MW 1, UTIL 1, CAL/MCDU 1, CMS 1
#
C H
B U S
A 2 2 A
FCM 3 A 2 GENERIC I/O 3 A 2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 NIC 5 (A) (ID = 33) CUSTOM I/O 2
A 2 A 2 A 2 A 2
SPARE SPARE FCM 4 A 2
B
U S
Power Supply 2 DC 2 ENGINE VIBE GPS 2 PSEM 2
AIOPB2
B
PROC 2 = CMF 2 U
#
S
A 1 2 B 2 B
A 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1
2 B 2 B 2 B
Power Supply 2 ESS 2/DC 2 BRAKES (INBD) CONTROL I/O 2 AIOPA2
PROC 1,4 - GENERATE CAS WARNINGS / MESSAGES 1 ACTIVE - 1 BACK-UP
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General Description The master warning red light comes on when an emergency condition occurs that requires immediate corrective flight crew action. The indicator light goes off when there is no longer a fail condition or when the master warning pushbutton annunciator is pushed. The master caution amber light comes on when an abnormal condition occurs that requires immediate crew awareness and subsequent corrective flight crew action. The indicator light goes off when there is no longer a fail condition or when the master caution pushbutton annunciator is pushed.
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Page 5
Figure 3: Central Warning System Schematic
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Notes:
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Page 7
Figure 4: EICAS Messages
TYPE
CAS MESSAGE APM FAIL APM FAULT APM MISCOMP AURAL WRN SYS FAIL AURAL WRN SYS FAULT AVNX ASCB FAULT AVNX MAU 1 FAN FAIL AVNX MAU 1 FAN FAULT AVNX MAU 1A FAIL
AVNX MAU 1B OVHT AVNX MAU 2 FAN FAIL AVNX MAU 2 FAN FAULT AVNX MAU 2A FAIL
CAUTION CAUTION ADVISORY CAUTION
AVNX MAU 2A FAULT
ADVISORY
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DESCRIPTION Failure of three or four Aircraft Personality Modules Failure of one or two Aircraft Personality Modules One or more APMs do not match The two aural warning channels are failed or are off One aural warning channel is failed Failure of at least one ASCB bus Three or four fans in MAU 1 failed One or two fans from MAU 1 have failed All functions hosted in MAU 1 channel A are unavailable
MAU 1 channel A has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy Modular Avionics Unit 1 A Overheat All functions hosted in MAU 1 channel B are unavailable MAU 1 channel B has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy Modular Avionics Unit 1B Overheat Two or three fans in MAU 2 failed Only one fan from MAU 2 has failed All functions hosted in MAU 2 channel A are unavailable MAU 2 channel A has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy
31400101
FOR TRAINING ONLY Reproduction Prohibited
31400601 31400501 31400201
31400701 31400802 31401602 31400402 31400102
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Notes:
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Page 9
Figure 5: EICAS Messages
AVNX MAU 2A FAULT
ADVISORY
AVNX MAU 2A CVHT AVNX MAU 2B FAIL
CAUTION CAUTION
AVNX MAU 2B FAULT
ADVISORY
AVNX MAU 2B OVHT AVNX MAU 3 FAN FAIL AVNX MAU 3 FAN FAULT AVNX MAU 3A FAIL
CAUTION CAUTION ADVISORY CAUTION
AVNX MAU 3A FAULT
ADVISORY
AVNX MAU 3A OVHT AVNX MAU 3B FAIL
CAUTION CAUTION
AVNX MAU 3B FAULT
ADVISORY
AVNX MAU 3B OVHT CCD 1 FAULT
CAUTION ADVISORY
Issue: June06 Revision: 00
MAU 2 channel A has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy Modular Avionics Unit 2A Overheat All functions hosted in MAU 2 channel B are unavailable MAU 2 channel B has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy Modular Avionics Unit 2B Overheat Two or three fans in MAU 3 failed Only one fan from MAU 3 has failed All functions hosted in MAU 3 channel A are unavailable MAU 3 channel A has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy Modular Avionics Unit 3A overheat All functions hosted in MAU 3 channel B are unavailable MAU 3 channel B has suffered failure conditions that does not affect the functionality, but may cause the loss of redundancy Modular Avionics Unit 3B overheat Failure of Pilot Cursor Control Device FOR TRAINING ONLY Reproduction Prohibited
31400102
31400602 31400502 31400202
31400702 31400803 31401603 31400403 31400103
31400603 31400503 31400203
31400703 31600131
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Notes:
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Chapter 31-52
Page 11
Figure 6: EICAS Messages
Issue: June06 Revision: 00
CAS MESSAGE CCD 2 FAULT CMS FAIL
TYPE ADVISORY CAUTION
CMS FAULT
ADVISORY
CREW WRN SYS FAULT CVR AFT FAIL
CAUTION ADVISORY
CVR FWD FAIL
ADVISORY
EICAS FAULT
CAUTION
EICAS OVHT FDR AFT FAIL
CAUTION ADVISORY
FDR FWD FAIL
ADVISORY
MFD 1 FAULT
CAUTION
MFD 1 OVHT MFD 2 FAULT
CAUTION CAUTION
MFD 2 OVHT NO TAKEOFF CONFIG PFD 1 FAULT
CAUTION WARNING CAUTION
PFD 1 OVHT PFD 2 FAULT
CAUTION CAUTION
PFD 2 OVHT SYS CONFIG FAIL
CAUTION CAUTION
VALIDATE CONFIG
CAUTION
DESCRIPTION Failure of Copilot CCD Failure of both instances of the Configuration Monitor System Failure of one instance of the Configuration Monitor System Failure of one or both Monitor Warning Functions Failure of the CVR Function of the Aft Digital Voice/ Data Recorder Failure of the CVR Function of the Forward Digital Voice/ Data Recorder The EICAS suffers failure conditions that may affect functionality The EICAS suffers an over temperature condition Failure of the FDR Function of the Aft Digital Voice/ Data Recorder Failure of the FDR Function of the Forward Digital Voice/ Data Recorder The MFD 1 suffers failure conditions that may affect functionality The MFD 1 suffers an over temperature conditions The MFD 2 suffers failure conditions that may affect functionality The MFD 2 suffers an over temperature condition The aircraft is not in a valid configuration for takeoff The PFD 1 suffers failure conditions that may affect functionality The PFD 1 suffers an over temperature condition The PFD 2 suffers failure conditions that may affect functionality The PFD 2 suffers an over temperature condition CMS detects a non-dispatchable configuration miscompare Miscompare of the Top Level System Part Numbers reported by the APM and the Part Numbers stored in the local non-volatile memory. Requires Part Number Configuration. FOR TRAINING ONLY Reproduction Prohibited
Indication and Recording System Diagnostic Tests SYSTEM DIAGNOSTICS MENU • Using the CCD No.2 touch pad to move the curser to the System Diagnostics Soft Key • Select the System Diagnostics Soft Key by pushing one of the enter keys on CCd No.2 • The System Diagnostics menu is displayed and a list of member systems organized by ATA chapter that have system diagnostic pages associated with them are presented.
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Chapter 31-52
Page 13
Figure 7: CMC Main Menu
MAINTENANCE M ESSA GES
SYSTEM DIA GNOSTICS
EXTENDED MAINTENA NCE
DATA LOADER
FILE TRANSFER
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Notes:
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Page 15
Figure 8: Indication and Recording Diagnostics
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170/190 MAINTENANCE TRAINING MANUAL
31-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 31-00 Digital Voice¦ Data Recorder ¦ (DVDR) System
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: ORIGINAL ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 12/16/2003 ¦ 31-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 31 INDICATING/RECORDING ¦ ¦ ¦ SYSTEMS ¦ ¦ ¦
¦ 21-00 Clock System
¦
¦
¦
¦
¦ ¦ ¦
1) Time Function on Digital Clock
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided at ¦ least one Time Function on MFD ¦ Status Page operates normally.
¦ ¦ ¦
¦ ¦ ¦
2) Time Function on MFD Status Page
C ¦ 2 ¦ ¦
¦ 1 ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
A ¦ 2 ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ May be inoperative provided: ¦ a) Time Function on Digital ¦ Clock operates normally, ¦ b) At least one Cockpit Voice ¦ Recorder (CVR) operates ¦ normally, and ¦ c) Repairs are made within 3 ¦ flight days.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
¦ ¦
¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ 22-00 Chronometer ¦ System
¦ ¦
¦ ¦ ¦
1) Chronometer Function on Digital Clock
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided at ¦ least one Chronometer Function on ¦ PFD operates normally.
¦ ¦ ¦
¦ ¦ ¦
2) Elapsed Time Function on Digital Clock
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
3) Chronometer Function on PFD
C ¦ 2 ¦ ¦
¦ 1 ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
C ¦ 2 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ Chronometer Function on Digital ¦ Clock operates normally.
¦ May be inoperative provided: ¦ a) At least one Flight Data ¦ Recorder (FDR) Function ¦ operates normally, and ¦ b) Repairs are made within ¦ three flight days.
¦ May be inoperative provided: ¦ a) At lease one Cockpit Voice ¦ Recorder (CVR) Function ¦ operates normally, ¦ b) Airplane is not dispatched ¦ from a designated airport as ¦ listed in the operator's MEL ¦ unless: ¦ 1 - The FDR failure occurs ¦ after pushback but prior to ¦ takeoff, or ¦ 2 - The FDR repair was ¦ attempted but was not ¦ successful. ¦ c) In those cases where repair ¦ is attempted but not ¦ successful, the aircraft may ¦ be dispatched on a flight or ¦ series of flights until the ¦ next designated airport ¦ where repair must be ¦ accomplished prior to ¦ dispatch, and ¦ d) Repairs are made within ¦ three flight days.
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 2 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 11/16/2004 ¦ 31-4 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 31 INDICATING/RECORDING ¦ ¦ ¦ SYSTEMS ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
2) Flight Data C ¦ 2 Recorder (FDR) ¦ Functions ¦ (Cont'd) ¦
¦ 1 ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
a) DFDR A ¦ Recording ¦ Parameters ¦ Required by ¦ Local ¦ Regulations ¦
¦ ¦ ¦ ¦ ¦ ¦
¦ May be inoperative provided: ¦ a) At lease one Cockpit Voice ¦ Recorder (CVR) operates ¦ normally, and ¦ b) Repairs are made within 20 ¦ calendar days.
¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
b) DFDR A ¦ Recording ¦ Parameters ¦ Not ¦ Required by ¦ Local ¦ Regulations ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ May be inoperative provided repairs ¦ are made prior to the completion of ¦ the next heavy maintenance visit. ¦ ¦ ¦ ¦
JOINT AVIATION AUTHORITIES MASTER MINIMUM EQUIPMENT LIST SUPPLEMENT AIRCRAFT EMBRAER 170 (1) System & Sequence Numbers Item
REVISION NO: DATE:
ORIGINAL 20 February 2004
PAGE S31-1
(2)Rectification Interval (3) Number installed (4) Number required for dispatch (5) Remarks or Exceptions
31 Indicating / Recording Systems -21-00 Clock System 1) Time Function on Digital Clock
C
1
0
2) Time Function on MFD Status Page
C
2
1
A
2
0
May be inoperative provided at least one Time Function on MFD Status Page operates normally.
May be inoperative provided: (a) Time Function on Digital Clock operates normally, (b) At least one Cockpit Voice Recorder (CVR) function operates normally, (c) The aeroplane does not exceed 8 further consecutive flights, and (d) Not more than 72 hours have elapsed since the Time Function on the MFD Status Page was found to be inoperative.
-31-00 Digital Voice Data Recorder (DVDR) System
last update: Jun06
-
-
-
As required by Operating Requirements.
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ATA 34 Navigation
170/190 MAINTENANCE TRAINING MANUAL
Table of Content
34-26 The Inertial Reference System (IRS) The Inertial reference system . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1
34-11 Integrated Electronic Stand-by System General Description
The dimming control bezel and screen lighting is provided by the “COMP/ CLK” potentiometer located on the right lighting control panel, although the screen lighting has a final adjustment through the IES bezel push button + and -.
The IES (Integrated Electronic stand-by) System is a back-up navigation source. The IES system is a solid-state, self-contained unit, which contains attitude sensors, a high-definition AMLCD (Active Matrix Liquid-Crystal-Display) associated with a back lightning box, electronic modular PCBs and a bezel for adequate man-machine interface. The equipment provides stand-by indication of: • Attitude • Airspeed • Altitude (standard or BARO (Barometric Setting) corrected and the associated barometric pressure adjustment) • MN (Mach Number) • VMO (Maximum Operating Velocity)/Mmo (maximum Mach Operation) and Vfe. • Skid/slip information • VS (Vertical Speed) • ILS (Instrument Landing System) deviations (from aircraft radio system ILS 1) A photocell, located on the right side of the IES units, adjusts automatically the brightness according to the cockpit lighting. The DC ESS BUS 1 and DC ESS BUS 2 supplies 28 VDC (Volt Direct Current) to the IES through the “INTEGRATED STBY PWR 1” and “INTEGRATED STBY PWR 2” 5A circuit breakers.
last update: Jun06
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Figure 1: IES Integrated Electronic Standby
Attitude
ILS ILS1 Slip/Skid information
STD 1013 hPa
VMO/MMO
3907M
340
Vfe
320 Airspeed
40
10
13 000 00 125 80 60
1 29 0 280
10
12 000
260 Mach Number
20
Vertical Speed
M. 45
CAGE last update: Jun06
Barometrically Corrected Altitude
EIS
BARO
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Components INTEGRATED ELECTRONIC STAND BY SYSTEM (IES) The IES is located on the cockpit main instrument panel, on the pilot side (left side). INTERFACE The IES interfaces with the generic I/O (Input/Output) #2 in MAU (Modular Avionics Unit) #2 through a discrete connection for the purpose of receiving a “turn on/off” altitude in meters. The IES presents the following interfaces: • Two ARINC (Aeronautical Radio Incorporated) 429 buses:one low-speed bus for anemobarometric data, and one with the generic I/O # 2 in MAU # 2 for the purposes of receiving radio information (ILS) and aircraft status for VMO/Mmo computation. • Five open/ground discrete inputs: four for aircraft configuration (instrument panel tilt angle, VMO/Mmo law), two for mode selection (normal operation/down loading), one for function selection (barometric pressure unit selection) and one for parity. • One discrete output for maintenance purposes. • One RS-422 digital bus for maintenance downloading. • Power input (28 VDC).
last update: Jun06
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Figure 2: Integrated Electronic Standby System
The aircraft’s Standby System has the following main components: EIS tegrated Electronic Standby System (IES) tificial horizon Altimeter Airspeed information
etic Standby Compass
These units can be used to navigate if the primary flight data systems are not functioning or are not available.
IES Indications ATTITUDE After an initialization of 90 s, and converting the sensor data to digital format, the IES computes attitude, taking in account the instrument panel tilt angle modelization coefficient, and the internal inertial sensor misalignment. The CAGE push button resets the attitude function to zero when pressed for more than 2 s. ALTITUDE Using the data from the integrated pitot/static/AOA sensor, the IES computes the standard altitude. The BARO corrected altitude is also computed, as a function of the barometric pressure, adjustment (BARO, in inHg or hPa, depending on the discrete input programming). The IES uses the pulses received from the BARO rotary knob to increase or decrease the barometric pressure value. Pressing the STD push button resets the BARO to the standard pressure. VERTICAL SPEED The VS data is given at the bottom of the altitude tape by a vertical green arrow oriented up or down and four green digits for the value. The arrow and the digits are in white rectangular window drawn with a single line. AIRSPEED Using the data provided by the integrated pitot/static/AOA (Angle of Attack) sensors, the IES computes the indicated airspeed. The airspeed range is from 30 kts to 520 kts, and the maximum range is limited below 520 kts if MN is > 1.
MACH NUMBER The IES computes the MN using the static and total pressure used for the altitude and airspeed processing. SKID/SLIP INDICATOR (LATERAL ACCELERATION) Using the lateral acceleration computed for the altitude processing, the IES displays the SKID/SLIP indication. The skid/slip data are given by a trapezoidal pointer moving, normally over the pitch scale, below the roll pointer. The pointer is white with a black background. The lateral acceleration (skid/ slip) symbol is a trapezoidal pointer displayed in black, surrounded in white, just below the roll pointer. VMO/MMO AND Vfe The IES computes the VMO/Mmo from the applicable VMO/Mmo law and computes the Vfe from the parameters with flap and slat extended. ABNORMAL OPERATION If the 28 VDC power is interrupted for more than 50 ms, the IES switches to low power state. During this state the computation is operational, but the display is stopped. When power returns the IES recovers the normal mode (if the interruption was less than 200 ms). In case of failure detection with a loss of information integrity, the IES enters the FAIL state and an OUT of ORDER page is displayed. The IES enters the ERROR state when a failure is detected about one or several functions, and the corresponding flag is displayed.
BAROMETRIC PRESSURE The barometric pressure range is 740 hPa to 1100 hPa (21.85 inHg to 32.48 inHg), with a resolution of 1 hPa (0.01 inHg) for the display and 0.25 hPa (0.01 inHg) for the display and 0.25 hPa (0.01 inHg) for the computation.
last update: Jun06
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Figure 3: IES Schematic
G/S LOC Arinc 429
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34-15 Air Data System (ADS) Components ADS consists of the following components: • • • •
4 Air Data Smart probes (ADSP) 2 Total Air Temperature (TAT) probes Guidance Panel - Baro Settings MAU Hardware: – I/O Modules – Three Air Data Application (ADA) Software on Processor Modules
Figure 1: ADS System Components
Display
Display BARO - SETTING KNOBS
Controller
ADSP 4
Controller
ADSP 3
TAT 2
TAT 1
ADSP 2
ADSP 1
MAU - 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B 16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9 8 2 B 7 6 5 4 3 2 1 #
B U S
C H
Power Supply 3 DC 1 AGM 1 CMC GPS 1 Power Supply 2 ESS 1 FCM 1 CUS TOM I/O 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1 AIOPB1 PROC 1 NIC 1 (A) (ID = 1) FCM 2
C H
B U S
last update: Nov06
MAU - 3
MAU - 2 #
A1 A1
A1 A1
16 15 14 13 12 11 10 9
A1 A1
A1 CONTROL I/O 1 A 1 BRAKES (OUTB D) A 1 PSEM 1 A1 A1 AIOPA1 Power Supply 1 ESS 1
ADA - AIR DATA APPLICATIONS
C H
B U S
2B 2B 2B
Power Supply 2 C H ESS 2/DC 2 BRAKES (INB D) CONTROL I/O 2 AIOPA2
2B
SPARE SPARE GENERIC I/O 2
B U S
C H
2B 2B
8 7 2B 6 5 2B 4 3 2 1 #
B U S
C H
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DA TABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2 Power Supply 1 DC 2
B U S
A1 A1 A1 A1 A1 A1 A
C H
B U S
#
B U S
16 1 15 14 13 12 1 11 10 1 9 1 1
C H
B
B B B B
8 1B 7 6 5 4 3 2 1B 1 #
B U S
C H
Power Supply 2 C BU H S DC 2 ENGINE VIBE GPS 2 A2 2 PSEM 2 A FCM 3 A2 GENERIC I/O 3 A2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 A2 NIC 5 (A) (ID = 33) A 2 CUS TOM I/O 2 A2 A2 AIOPB2
ADC
SPARE SPARE FCM 4
A2 Power Supply 1 C BU H S ESS 2
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Electronic Displays System (EDS) The altitude, speed and other data are transmitted to the EDS via ASCB bus, and displayed on the Primary Flight Displays (PFD). Aircraft Altitude is displayed on the right side of the attitude display. Airspeed information is displayed on the left side of the attitude display.
Figure 2: Electronic displays System (EDS)
ADS provides data for: - altimeter - airpeed - AOA - mach - pressure - TAT - SAT
STATIC AIR TEMP TOTAL AIR TEMP TRUE AIR SPEED
Altitude in Ft Altitude in M Vertical Speed (FPM) Airspeed
Baro Correction (InHg or HPA)
Mach
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Air Data Systems • ADS1 is the Air Data System 1 for the Pilot and Flight Control System (FCS) source 1. Static Pressure (Ps) is corrected for beta effects by averaging the local static pressure values from ADSP1A and ADSP2A. ADSP1A then applies a Mach and AOA correction on the beta corrected Ps to achieve a True Ps value. • ADS2 is the Air Data System 2 for the Co-Pilot and Flight Control System (FCS) source 2. Static Pressure (Ps) is corrected for beta effects by averaging the local static pressure values from ADSP3A and ADSP4A. ADSP4A then applies a Mach and AOA correction on the beta corrected Ps to achieve a True Ps value. • ADS3 is the Air Data System 3 for the stand-by and Flight Control System (FCS) source 3. Static Pressure (Ps) is corrected for beta effects by averaging the local static pressure values from ADSP3B and ADSP4B. ADSP3B then applies a Mach correction on the beta corrected Ps to achieve a mach and beta corrected Ps value. • ADS FC is the source 4 of air data pressure for Flight Control System (FCS). Static Pressure (Ps) is corrected for beta effects by averaging the local static pressure values from ADSP1B and ADSP2B. ADSP2B then applies a Mach correction on the beta corrected Ps to achieve a mach and beta corrected Ps value. • Air Data System 3 and the Stand-by display are capable of utilize local Ps in the event of averaged Ps failure. A CAS messages will alert the pilot about the Loss of Side Slip Compensation.
RVSM The ADS 1 and 2 are RVSM (Reduced Vertical Seperation Minimum) compliant. ADS 3 is RVSM compliant when operating in normal mode (not in SLIPCOMP fail mode). The surrounding area around each integrated pilot/ static/ AOA (Angel of Attack) sensor must be smooth since it can impact the RVSM performance. Procedures can be found in the aircraft maintenance manual to do the inspection/ check of the RVSM critical region, and the inspection check of the RVSM critical components (including the integrated pilot/ static/ AOA sensors and radome.
RVSM General The purpose RVSM is to increase the number of aircraft safely occupying the same volume of air space. RVSM reduces the required vertical seperation of aircraft from 2000‘to 1000‘without diminishing or impairing flight safety. this allows more aircraft to use popular and previously congested air space and routes. This is safely accomplished by increasing the accuracy, sensitivity and reliability of aircraft instruments, most importantly the altimeter. RVSM only applies to altitudes between 29000‘ and 41000‘ (FL290 and FL410), below 29000‘ a vertical separation of 2000‘ is the standard. To fly in RVSM airspace aircraft must be approved for RVSM operations (carry special equipment and be certified). An exception from this is made for non-equipped State aircraft, e.g. military fighters.
Figure 3: Air Data Systems
AIR DATA SYSTEM 2
AIR DATA SYSTEM 3
ADA 2
MAU 2
MAU 3
ADA 3
TAT 2
4A 4B
3A ADSP 4
ADSP 3
TAT 2
3B TAT 1
TAT 1
ADSP 2
4A 4B
ADSP 1
3A ADSP 4
1A
2A
TAT 2
ADSP 1
1A
2A
AIR DATA SYSTEM 1 MAU 1
2B
3B
TAT 1
ADSP 2
1B
2B
ADSP 3
AIR DATA SYSTEM FC
1B
ADS Displays, Normal Conditions
ADA 1 AIR DATA SYSTEM 1
AIR DATA SYSTEM 2
AIR DATA SYSTEM 4 3
ARINC 429
(IES) INTEGRATED ELECTRONIC STANDBY
No 1 PFD
last update: Nov06
No 2 PFD
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Notes:
Figure 4: ADC System - Block Diagram DC BUS 1
DC BUS 1
PROBE 1A-2A
PROBE 1B-2B
5
5
FWD FUSELAGE
FWD FUSELAGE
INTEGRATED PITOT/STATIC/ADA SENSOR NO. 1
INTEGRATED PITOT/STATIC/ADA SENSOR NO. 2
CHANNEL A (ADS NO. 1)
CHANNEL A (ADS NO. 1)
CHANNEL B
CHANNEL B
ARINC 429 ARINC 429 TAT NO. 1
ARINC 429 ARINC 429
FWD FUSELAGE FWD AVIONICS CMPT
MAU NO. 1 GENERIC I/O MODULE
PROC NO. 1 (ADA FUNCTION)
CUSTOM I/O MODULE
NIC
ASCB
TO OTHER AIRCRAFT AVIONICS SYSTEMS
BACKPLANE
ADC SYSTEM - BLOCK DIAGRAM
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Figure 5: Notes: Air Data System Architecture
Component
ADS 1
ADS 2
ADS 3
IES
ADS FC
ADSP
ADSP 1 CH A ADSP 2 CH A
ADSP 3 CH A ADSP 4 CH A
ADSP 3 CH B ADSP 4 CH B
ADSP 3 CH B ADSP 4 CH B
ADSP 1 CH B ADSP 2 CH B
ADA
ADA 1
ADA 2
ADA 3
-
-
MAU
MAU 1
MAU 2
MAU 3
-
MAU 1
MAU I/ O
GEN I/ O 1
GEN I/ O 2
GEN I/O 3
-
CUS I/O 1
ASCB BUS
ASCB 1
ASCB 2
ASCB 2
-
ASCB 1/2
AC POWER
DC 1/ ESS 1
ESS 3/ DC 2
DC 1/ ESS 2
ESS 2
DC 1
TAT
TAT 1
TAT 2
TAT 1
-
-
Figure 6: ADS System - Block Diagram RH CBP COCKPIT DC ESS BUS 3
Integrated Pitot/Static/AOA Sensor (ADSP) The integrated pitot/static/AOA sensor through of the pressure transducers and differential pressure transducers sense the pressures (pitot, static and differential) and through the algorithms implemented in the software compute measured values known as the local total pressure (PtL), local static pressure (PsL), and local angle of attack (AOAL). These local air data values and TAT (Total Air Temperature) through the TAT sensor are sent to the air data application (ADA), which is located in the MAU (Modular Avionics Unit)s. The IES (Integrated Electronic Stand-by) receives the air data parameters of the integrated pitot/static/AOA sensor. The integrated pitot/static/AOA sensor combines pitot, static, angle of attack, and air data computer functions into one LRU providing air data parameters to the air data system flight controls system and stall warning and protection system. The following functions are performed by each channel A of the integrated pitot/static/AOA sensors: • Pitot pressure sensing • Static pressure sensing • Angle of attack sensing • TAT sensing (when TAT connected) • Failure annunciation • Current monitoring for MFP and TAT sensors (when TAT connected) • MFP and TAT sensor heat control
• Perform SSEC on local static pressures The following functions are performed by each channel B of the integrated pitot/static/AOA: • Pitot pressure sensing • Static pressure sensing • Failure annunciation • Monitoring of MFP and channel A controller • Channel A MFP heater controller bypass • Perform SSEC on local static pressures
Figure 7: Multi-Function Probe (MFP)
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Multi function Probe (continued) Local Angle of Attack The local AOA data is received by the Generic I/O cards on ARINC 429 from the ADSPs channel A and made available on the ASCB bus to the Stall Protection Computer and Windshear Guidance Computer. Local AOA data out putted by ADSPs is not considered part of the ADS system. The AOA information is utilized on the ADS system only during the Ps SSEC correction performed internally on the ADSPs channel A. Barometric Correction The baro correction input is a two bit code read off two open ground input discretes from the baro knob. The I/O function will interface with the knob input and transmit the raw data to the ADA function. The knob has 16 discrete positions, or clicks, per revolution, broken into 4 2-bit codes transmitted by the discretes. Total Air Temperature The TAT input to the ADS will be from the ADSP. The I/O function will receive the data via an ARINC 429 interface and will communicate the information to the ADA function. The measurement of TAT is accomplished by two TAT probes projecting into the airstream. The TAT contains a resistive element that changes resistance with a change in temperature. Discrete Inputs The ADA Interfaces directly with aircraft discretes that provide information such as weight-on-wheels (WOW) indication, baro standard setting, TAT heater status (from the ADSPs),etc. Some of these interfaces are not specific to the ADS and will not be directly controlled by the ADS. Instead the discretes will be read by the I/O function and relayed to the ADA via the Virtual Back plane. The ADA will utilize data from the I/O function on an as-needed basis, based on the application.
Figure 8: ADSP Overview
P PT
P
PS
PS
PS
PT
PS
TRANSDUCER
TRANSDUCER
TRANSDUCER
P
P
TRANSDUCER
TRANSDUCER
ELECTRONIC CIRCUITS CH A
CH B
LEGEND : PNEUMATIC SIGNAL ELECTRICAL SIGNAL
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Air Data Computer (ADC) The Air Data Computer is physically and pneumatically connected to the MFP. The ADC performs air data (pressure) calculations, which are forwarded to the Honeywell Air Data Application (ADA), resident in a Modular Avionics Unit (MAU), or to the Sextant Stand-by Indicator.
Circuit Card Assemblies (CCA) Differential Pressure CCA The differential pressure card contains the analog signal conditioning and analog to digital conversion circuitry that interfaces the differential pressure sensor to the system. This card also contains the Programmable Logic Device (PLD) that is the interface between the differential pressure circuits and Honeywell APM to the CPU CCA. Air Pressure Module (APM) CCA The Honeywell APM card contains the analog signal conditioning and analog digital conversion circuitry that interfaces the absolute pressure sensors to the system. The APM also provides the temperature of the sensors in a digital format. Central Processing Unit (CPU) CCA The micro controller card directs the data acquisition of measured pressures and temperatures, ARINC communications, the calculation of air data parameters, air data warnings, and resident BIT capabilities with NVRAM fault storage. Power Supply CCA The power supply CCA converts the 28 VDC input to useable voltages required by the internal circuitry. To protect the unit from switching transients and effects from lightning strikes, a complete transorb circuit is used. The power supply CCA also incorporates reverse polarity protection.
Figure 9: New ADSP Card
Lightning EMI Power B Power A CPU B CPU A dP and PCADS Conditioning HI APM New PCADS
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Components AIR DATA APPLICATION FUNCTION The ADA function is the primary component in the ADC system. The ADA function receives the air temperature, pressure, and AOA data from the sensors. The ADA function receives inputs from the other systems or applications in the aircraft that are used to calculate and correct the air data parameters. The ADA function makes the necessary corrections before it sends the air data parameters to other avionics systems. There are three ADAs found in the MAUs. They are hosted in processor modules on a DEOS (Digital Engine Operating-System) platform in the MAUs. The ADA 1 is on the PROC 1 in the MAU 1, ADA 2 is on PROC 4 in the MAU 2, and ADA 3 is on the PROC 5 in the MAU 3. The ADA function receives the air data parameters from integrated pitot/static/AOA sensor through the generic I/O (Input/output) module in the MAU. It also receives the barometric correction inputs from a baro set control on the GP (Guidance Panel). The ADA function then takes these inputs and calculates the air data parameters and correction signals. The ADA function sends the air data through the MAU backplane to the NIC. The NIC transmits the air data on the ASCB (Avionics Standard-Communication Bus) to the other avionics systems. The IES unit can shows the air data parameters even if the ADA function has a fault. The IES unit receives air data parameters directly from integrated pitot/static/AOA sensor 3. The Integrated/Pitot/Static/AOA sensor make the corrections for the probe type, probe position, and the airflow errors where the errors are aircraft type or model dependent as follow: • Static source error correction (SSEC) is a set of constants used to correct the static pressure. • Maximum operating airspeed/mach.
The outputs of the ADA function are the air data parameters that follows: • • • • • •
Angle of attack Corrected static pressure Corrected total pressure Impact pressure Dynamic pressure Total air temperature
TAT Sensor The forward fuselage of the aircraft has two TAT sensors that sense the TAT of the external air. The sensor reads the air temperature with a wire probe that projects into the airstream. Each sensor contains a 500 ohm resistor that changes the value with the changes in the temperature. The sensor housing has a hermetic seal that prevents the contamination of the sensor. The TAT sensor sends the Integrated/Pitot/Static/AOA an analog signal, which shows as degrees (°C) on the cockpit displays. Each TAT sensor has a three-wire direct connection to the Integrated/Pitot/ AOA sensor. The Integrated/Pitot/AOA sensor receives the TAT data through the three-wire direct connection and then sends the data to the generic I/O module in the related MAU. The generic I/O module sends the TAT data through the backplane to the ADA processor module, which uses the TAT data to calculate the air temperature. The Integrated/Pitot/Static/AOA sensor does a TAT compensation algorithm to adjust the measured temperature for the aircraft wiring resistance, mach recovery factor, and probe heater effects. Each TAT sensor uses a 28 VDC (Volt Direct Current) heater to keep the sensor free from ice. This lets the sensor supply accurate data in all flight conditions. The Integrated/Pitot/Static/AOA sensor supplies the power to the TAT heater. The three-wire connection lets the Integrated/Pitot/Static/AOA sensor monitor the TAT sensor for the ice conditions. The Integrated/Pitot/ Static/AOA sensor adjusts the power to control the heater current as necessary to keep the TAT sensor free from ice.
Figure 11: TAT Probes
Total Air Temperature Probe
Generic I/O
ARINC 429
ADA last update: Nov06
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Ice Protection for Smart probe and TAT Each of the Smart Probes and TAT’s include resistive heater elements to accomplish de-icing and anti-icing capabilities of the probes for continued sensor accuracy in icing environments. Each Smart Probe is capable of monitoring and controlling the power applied to its own MFP as well as a TAT probe, if a TAT probe interfaces with the Smart Probe. The probe heat control functionality significantly increases the life of the MFP and TAT by providing the necessary power required to keep the probe free of ice. Therefore, when the Smart Probe’s and TAT’s are functioning during aircraft ground operations, minimal heat is required and applied to the probes as opposed to conventional probes which will continue to have maximum power applied. In cold temperature environments, however, the MFP and TAT will be provided the additional power necessary to maintain ice-free operation. In AUTO mode the A/C electrical system will apply power to the MFP and TAT heaters if the aircraft is airborne (WOW False) or if any Engine is running.
Figure 12: Ice protection for Smart probe and TAT
SMART PROBE MONITORS AND CONTROLS THE POWER FOR HEATING IT`S OWN PROBE AS WELL AS A TAT PROBE.
ENG1 running Heater power ON
ENG2 running
WOW WOW False
commands relay to OFF = Heater power ON commands relay to ON = Heater power OFF
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Baro Set Controls A baro set control on the GP lets the pilot and copilot select the barometric correction either in inches of mercury or hectopascals. The GP is installed on the instrument panel in the cockpit. The set barometric correction value is shown under the altitude tape on the PFD. The baro set control supplies the baro correction inputs to the ADS. There is one baro set control for each ADS for use by the pilot and copilot. The barometric correction input is a 2-bit code read from the two open/ ground discretes from the baro set control. The baro set control has 16 open/ ground discrete positions per revolution that are divided into the four 2-bit codes. The position of the baro set control sets the baro correction value that is sent to the ADS. The control I/O module in the MAU connects with the baro set control through an RS-422 bus to receive the barometric correction input discretes (A and B) from the baro set control. The control I/O module changes these input discretes into a 7-bit count value that is related to the position of the baro set control (number of clicks). The control I/O module then sends this 7-bit barometric correction code through the MAU backplane to the ADA processor module. The ADA processor module uses this barometric correction value to calculate the air data parameters. There are two independent STD push buttons on the baro set control knobs on the guidance panel. Each switch can be pushed to set a standard baro correction value for the applicable ADS. The normal barometric correction values are the 29.92 inHg and the 1013 mbarMillibar. The status of the STD switch is also sent through the RS-422 bus to the control I/O module, and then through the backplane to the ADA processor module.
IN/HPA switch selects baro correction format: inches of mercury or HectoPascals. PUSH STD resets correction to standard value (29.92 inHg or 1013hPA). last update: Nov06
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Abnormal Operation With a triplex Air Data System installation and no failures, the Air Data readouts in normal operation on the pilot’s and copilot’s PFDs are from ADS 1 and ADS 2, respectively. Upon failure of either ADS 1 or ADS 2, only the effected side will lose the ADS readouts and a failure indication will be provided to the flight crew. The effected side will automatically revert or pilot can source select (using the Reversionary Panel) one of the remaining two ADS’s for the associated PFDs air data display. The reversion operation provides the pilots with multiple ADS availability. For 1 or 2 ADS failures, the reversionary logic for the pilot and copilot PFDs are as follows: Pilot PFD - Reversionary Logic • ADS 1 (Normal Operation) • ADS 3 (1st Reversion) • ADS 2 (2nd Reversion) Copilot PFD - Reversionary Logic • ADS 2 (Normal Operation) • ADS 3 (1st Reversion) • ADS 1 (2nd Reversion) The automatic reversion will occur when: the currently displayed Air Data source id indicated as invalid on ASCB - OR - the currently displayed airspeed is indicated as invalid on ASCB - OR - the currently displayed altitude is indicated as invalid on ASCB. If no valid source of air data information is available the source shall return to the default for the given PFD (i.e. ADS 1 for pilot and ADS 2 for copilot). The automatic reversion shall not return to the “correct” source should that source become valid subsequent to the reversion. After manual or automatic reversion, the selected source flag is displayed on the reverted side and reversion button (on the reversionary panel) is illuminated.
Revert ADS1 as necessary if AUTO reversion did not act. Reference checklist to determine dispatch or continued mission impact. Revert ADS2 as necessary if AUTO reversion did not act. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: ADS3 not available for reversion. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: Dedicated ADS to Flight Control Failed. Reference checklist to determine dispatch or continued mission impact. Revert ADS1 as necessary or avoid Ice-conditions. Reference checklist to determine dispatch or continued mission impact. Revert ADS2 as necessary or avoid Ice-conditions. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: ADS3 not reliable for reversion or avoid Iceconditions. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: ADS4 not reliable for reversion or avoid Iceconditions. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: Reduced HTR availability for ADS1. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: Reduced HTR availability for ADS2. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: TAT/SAT/TAS 1and 3 failed or affected by ice conditions. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: TAT/SAT/TAS 2 failed or affected by ice conditions. Reference checklist to determine dispatch or continued mission impact. Crew Awareness: Possible altitude errors for ADS3 if flying in side slip conditions.
-
Crew Awareness: Possible altitude errors for ADS4 if flying in side slip conditions.
14
ADS 4 SLIPCOMP FAIL*
ADVISORY
Action
Figure 15: ADS Fault Codes - FIM Fault Code Index
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Notes:
Figure 16: Maintenance Messages
TASK 34-15-00-810-825-A Failure of Integrated pitot/static/AOA sensor 1 A. General (1)
This task is for fault code 34130001UT1 (AIR/AOA SNSR1 FAULT [UTIL]).
B. Fault Description (1)
The utilities 1 senses a failure of the integrated pitot/static/AOA sensor 1 and sends a fault report to the CMC.
C. Probable Causes (1)
Failure of the integrated pitot/static/AOA sensor 1 (AMM MPP 34-13-01/401) (AIPC 34-13-01).
(2)
Defective harness (WM 34-13-50).
D. Circuit Breaker List TYPE
DESIGNATION
LOCATION
BUS
LOCATION TIP
MCDU PAGE
CB
ADS 1 PROBE 1A-2A
LHCBP
DC BUS 1
NAV
-
CB
ADS FC PROBE 1B-2B
LHCBP
DC BUS 1
NAV
-
CB
MAU 1 PWR 1
LHCBP
DC ESS BUS 1
-
-
CB
MAU 1 PWR 2
LHCBP
DC ESS BUS 1
-
-
CB
MAU 1 PWR 3
LHCBP
DC BUS 1
-
-
E. Fault Isolation Procedure
SUBTASK 34-15-00-862-045-A CAUTION: MAKE SURE THAT THE INTEGRATED PITOT/STATIC/AOA SENSORS, TAT SENSORS, ICE DETECTORS AND STATIC PORT HAVE NO COVERS ON THEM BEFORE YOU DO THE MAINTENANCE PROCEDURE. THESE COMPONENTS CAN BECOME HOT DURING THE MAINTENANCE PROCEDURE. AS A RESULT, DAMAGE TO THEM WILL OCCUR IF YOU DO NOT REMOVE THE COVERS. (1)
Open these circuit breakers and, after approximately 5 seconds, close them: NOTE: If the power is removed from the CMCM for more than 10 seconds, the CMCM will start its shutdown procedure. When the power is supplied to the CMCM again, it will start its power-up procedure. The full process takes approximately 5 minutes:
· ·
last update: Nov06
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MAU 1 PWR 1-LHCBP (DC ESS BUS 1) MAU 1 PWR 2-LHCBP (DC ESS BUS 1)
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ADS System Diagnostics Tests SYSTEM DIAGNOSTICS MENU Select System Diagnostics in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the System Diagnostics Soft Key • Select the System Diagnostics Soft Key by pushing one of the enter keys on CCD No.2. • The System Diagnostics menu is displayed and a list of member systems organized by ATA chapter that have system diagnostic pages associated with them are presented.
Figure 17: CMC Main Menu
MAINTENANCE M ESSA GES
SYSTEM DIA GNOSTICS
EXTENDED MAINTENA NCE
DATA LOADER
FILE TRANSFER
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Notes:
Figure 18: ADS System Diagnostic Test
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34-26 The Inertial Reference System (IRS) The Inertial reference system Each IRS system consists of a Micro IRU, Mounting Tray, and an Aircraft Personality Module (APM). The IRUs are installed in the forward compartment, as shown. Each Micro IRU is installed in a special mounting tray and held in position by two hold-down screws.
The Inertial Reference Unit (IRU) The main function of the IRU is to sense and compute linear accelerations and angular turning rates about the airplane`s pitch, roll and yaw axes. This data is used for pitch and roll displays and navigational computations. The IRU contains six sensors, three ring laser gyros that measure angular motion about the longitudinal, lateral and vertical axis, and three accelerometers that measure linear motion about the longitudinal, lateral and vertical axis.
Power supplies For the micro IRU 1, the ESS (Essential) Bus 1 supplies the primary power, and ESS Bus 2 supplies the secondary power.For the micro IRU 2 , the DC (Direct Current) Bus 2 supplies the primary power, and DC Bus 1 supplies the secondary power.This electrical configuration isolates the two micro IRUs and prevents failures across the different power buses.
Figure 1: IRU, Configuration Module, and Tray
IRU, Configuration Module, and Tray
IRS components are located in the forward electronics bay.
last update: Nov06
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Electronic Displays System (EDS) The attitude information is transmitted from the Generic I/O via ASCB-D bus to the EDS, and is displayed to the crew on the PFDs. Attitude Display The Attitude Display (ADI) consists of an artificial horizon, pitch tape, roll pointer/scale, and aircraft symbol. The ADI will extend from the airspeed tape to the altitude tape with a truncated sphere outline. The artificial horizon consists of a horizon stabilized sky/ground solid colour representation. A reference aircraft symbol is displayed in the ADI. Pitch Tape A linear pitch tape is displayed through the centre of the attitude display. The pitch tape is horizon stabilized. The pitch tape appears to go behind the aircraft symbol and the radio altitude digits whenever visible. Roll Scale A linear roll scale is displayed on the top of the truncated ADI sphere. Primary source for Pilot PFD is IRS1, for Copilot is IRS2. Source selection (reversion) is available pressing IRS button on the reversionary panel. Button is illuminated to indicate that no primary source is selected and the PFD indicates the selected source, other than the primary.
Figure 2: Electronic Display System (EDS)
IRS SYSTEM PROVIDES: - Primary aircraft attitude Pitch angle Roll angle - Magnetic and true heading - Inertial velocity - Navigation position - Wind data last update: Nov06
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34-26-03 IRS Operation and Indication Position Initialization
MCDU Position Initialization Page
The IR component requires system initialization (entry of latitude and longitude). Initialization may take place either from a Flight Management System (FMS) from input that the crew manually enters via the Multi-Functional Control Display Unit (MCDU), or automatically from the Global Positioning System (GPS). A pilot-entered position has priority over a position from a GPS.
To initialize the FMS position, select the POS INIT line select switch prompt from either the NAV IDENT or POS SENSORS page. The POS INIT page will list positions that can be line-selected for initialization of the FMS using the LOAD prompt. Any of the positions listed can be used for the initialization, or the pilot may enter the appropriate Latitude/Longitude or reference way point using prompt 2L.
The MCDU page allows the pilot to manually initialize the position sensors. The position entered becomes the FMS position and is the position given to the long range sensors.
In addition, the IR component receives Air Data information such as altitude, altitude rate and true air speed from an Air Data System (ADS). The Micro IRU gathers this information and produces the parameters for Body Frame, Local Level Frame, and Earth Frame.
Figure 1: Position Initialisation
IRU units Position Initialization Pilot manual input GPS
last update: Dec06
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Reversion Primary source for Pilot PFD is IRS1, for Copilot is IRS2. Source selection (reversion) is available pressing IRS button on the reversionary panel. Button is illuminated to indicate that no primary source is selected and the PFD indicates the selected source, other than the primary.
Figure 2: Reversionary Mode PILOT PFD
IRS2
CO - PILOT PFD
IRS2
REVERSIONARYPANEL DISPLAYS AUTO PFD
MFD
SENSORS ADS IRS
EICAS MFD MODE
If reversionary mode is selected: – – – –
FD modes disengage AP modes disengage FD/AP can be re- engaged Annunciates: - IRU1 - IRU2
last update: Dec06
IRU 1
FOR TRAINING ONLY - Reproduction Prohibited
IRU 2
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Micro Inertial Reference System Aircraft Personality Module (APM) The Micro IRS APM is programmed to contain configuration and installation data. The APM facilitates Original Equipment Manufacturer selectable features using common IRU Part Numbers. It also facilitates storage of Configuration Selects and Tray Alignment Euler Angles electronically. Information contained in the APM file includes: • • • • • • •
Aircraft Type/Serial Number/SDI Program MAGVAR Select Program Dedicated IRU Battery Select Mount Misalignment Euler Angles Output Filter Characteristics Other A/C Specific Data CRC
The APMs are initially programmed as a part of the tray alignment procedure. The Configuration may be changed later without modifying the Installation Specific data.
Figure 3: Aircraft Personality Module
APM is programmed to contain configuration and installation data. APM information includes: Aircraft Type/Serial Number Source Destination Identifier (SDI) Program MAGVAR Select Mount Misalignment Data Other A/C Specific Data
APMs are initially programmed as a part of the tray alignment procedure.
last update: Dec06
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Abnormal Operation / CAS messages The IRS has been designed such that in the event of a normal aircraft power interrupt or power transient, no degradation in performance will occur due to switching or operating from the backup power source. The two Micro IRUs receive power from different primary and different secondary power sources, so in the event of a loss of power to one Micro IRU, the other Micro IRU would not also experience a power loss at the same time. In case of mode failure, power loss or loss of one IRU, the affected station can source select the other side station by means of the IRS reversionary panel switch, and a CAS message will be displayed to alert the crew. The CAS messages, its priorities and the actions that should be taken by the crew are shown on the table below.
Figure 4: CAS Messages
FAULT CODE
FAULT DESCRIPTION
MAINTENANCE
GO TO FIM TASK
MESSAGE
34200101
34200102
34200201
34200202
IRS 1 FAIL (CAUTION)
IRS 2 FAIL (CAUTION)
IRS 1 NAV MODE FAIL (ADVISORY)
IRS 2 NAV MODE FAIL (ADVISORY)
34200300
IRS PRES POS INVALID (ADVISORY)
34200400
IRS ALIGNING (ADVISORY)
34200500
IRS EXCESSIVE MOTION (CAUTION)
last update: Dec06
34265515IR1
34 - 26 - 00 - 810 - 801 - A
34265519IR1
34 - 26 - 00 - 810 - 803 - A
34265529IR1
34 - 26 - 00 - 810 - 807 - A
34265515IR2
34 - 26 - 00 - 810 - 808 - A
34265519IR2
34 - 26 - 00 - 810 - 810 - A
34265529IR2
34 - 26 - 00 - 810 - 814 - A
34265515IR1
34 - 26 - 00 - 810 - 801 - A
34265519IR1
34 - 26 - 00 - 810 - 803 - A
34265529IR1
34 - 26 - 00 - 810 - 807 - A
34265515IR2
34 - 26 - 00 - 810 - 808 - A
34265519IR2
34 - 26 - 00 - 810 - 810 - A
34265529IR2
34 - 26 - 00 - 810 - 814 - A
34265515IR1
34 - 26 - 00 - 810 - 801 - A
34265515IR2
34 - 26 - 00 - 810 - 808 - A
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Notes:
Figure 5: Maintenance Messages
MAINTENANCE MESSAGE FAULT CODE
last update: Dec06
MAINTENANCE MESSAGE NAME
GO TO FIM TASK
34265515IR1
IRU1 APM/WRG FAULT
34 -26 - 00 - 810 - 801- A
34265515IR2
IRU2 APM/WRG FAULT
34 - 26 - 00 - 810 - 808 - A
34265519IR1
IRU1 FAULT
34 - 26 - 00 - 810 - 803 - A
34265519IR2
IRU2 FAULT
34 - 26 -- 00 - 810 - 803 - A
34265520IR1
[IRU1] LEFT GEN INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 804 - A
34265520IR2
[IRU2] RIGT GEN INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 811 - A
34265521IR1
[IRU1] RIGT GEN INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 815 - A
34265521IR2
[IRU2] LEFT GEN INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 818 - A
34265522IR1
[IRU1] GPS2 INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 806 - A
34265522IR2
[IRU2] GPS1 INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 812 - A
34265523IR1
[IRU1] GPS1 INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 805 - A
34265523IR2
[IRU2] GPS2 INPUT BUS/WRG FAULT
34 - 26 - 00 - 810 - 813 - A
34265529IR1
IRU1 PROGRAM PIN PARITY/WRG FAULT
34 - 26 - 00 - 810 - 807 - A
34265529IR2
IRU2 PROGRAM PIN PARITY/WRG FAULT
34 - 26 - 00 - 810 - 814 - A
34267016IR1
IRU1 SECONDARY POWER/WRG FAULT
34 - 26 - 00 - 810 - 816 - A
34267016IR2
IRU2 SECONDARY POWER/WRG FAULT
34 - 26 - 00 - 810 - 819 - A
34267017IR1
IRU1 PRIMARY POWER/WRG FAULT
34 - 26 - 00 - 810 - 817 - A
34267017IR2
IRU2 PRIMARY POWER/WRG FAULT
34 - 26 - 00 - 810 - 820 - A
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Notes:
Figure 6: CMC Navigation tests
CMC Navigation tests
last update: Dec06
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Notes:
Figure 7: CMC Navigation tests
CMC Navigation tests
last update: Dec06
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34-31 Radar Altimeter System General The Radar Altimeter System is a baseline system on board the ERJ 170 aircraft. A second Radar Altimeter System is optional. The ERJ 170 Radar Altimeter System provides the pilots with dependable, accurate altitude, up to + 2,500 feet above ground level (AGL) during the approach and landing phases of aircraft operation. The radar altitude is displayed on both pilot and copilot Primary Flight Displays (PFD’s).
last update: Nov06
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Chapter 34-31
Page 1
Figure 1: Radar Altimeter System
The Radar Altimeter System Provides:
ABSOLUTE ALTITUDE
TRUE ALTITUDE
PRESSURE ALTITUDE
29.921 IN. HG.
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LRU Each ERJ 170 Radar Altimeter System consists of a Receiver/Transmitter mounted in a tray, a Configuration Module, a Receive Antenna and a Transmit Antenna.
Installation The ERJ 170 Radar Altimeter System #1 Receiver/Transmitter (R/T) is located at the after cargo compartment beside the cargo door, on the LH side, the Configuration Module is located beside the R/T #1 and the two Antennas are located at the lower part of Centre Fuselage III. The optional ERJ 170 Radar Altimeter System #2 R/T is located at the after cargo compartment beside the cargo door, on the RH side, the second Configuration Module is located beside the R/T #2 and the second pair of Antennas are located at the lower part of Centre Fuselage III. The locations of the antennas insure acceptable operation at the normal extremes of pitch and roll.
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Chapter 34-31
Page 3
Figure 2: Component location
RAT 1 RAT 3 RAR 1 RAR 3 RAR 2 RAT 2
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Interfaces EGPWS Interface
Configuration Module Interface The R/T interfaces with the Configuration Module for storage of the zero feet offset.
TCAS Interface The ERJ 170 Radar Altimeter System interfaces with the Traffic Alert and Collision Avoidance System (TCAS) transmitting the Radar Altitude data via an ARINC 429 bus, the TCAS also receives a Radar Altitude Valid signal. The TCAS uses the radar altitude to inhibit descend resolution advisories.
MAU Interface Also the MAU receives the radar altitude from the R/T via an ARINC 429 bus at the Generic I/O Module. The Generic I/O Module contains a standard interface circuit that transfers data to and from a back plane bus. The interface circuit performs functions that include data distribution, data integrity checking, and source identification. The back plane bus is a parallel high capacity general-purpose bus contained in the MAU that transfers all data between the modules and the network interface controller. The network interface controller also contained in the MAU is a dedicated module that interfaces the back plane bus to the external ASCB network.
The ERJ 170 Radar Altitude System also interfaces with the Enhanced Ground Proximity Warning System (EGPWS) via the ASCB. The radar altitude and how fast the airplane sink or the ground profile change is a key parameter for the EGPWS. DVDR Interface The ERJ 170 Radar Altimeter System also interfaces with the Digital Voice Data Recorders (DVDR) via an ARINC 717 bus. Radar altitude is one of the mandatory parameters that must be recorded.
AMDS Interface The ERJ 170 Radar Altimeter System interfaces with the Aircraft Diagnostics and Maintenance System (ADMS) also contained in the MAU, which allows fault and troubleshooting information for the system to be displayed and retrieved by maintenance personnel.
The ASCB network provides the communication between the MAU and the DU`s, where the Radar Altitude will be displayed on the PFD’s
CENTER FUS III RADAR ALTIMETER 1 CONFIGURATION MODULE (SDS 34-31) (MPP 34-31-03)
Rad alt 2
SPDA 1 (SSM 24-61-80) CENTER FUS III RADAR ALTIMETER 2 CONFIGURATION MODULE (SDS 34-31) (MPP 34-31-03)
E
28 VDC
F
DC POWER - M5 DC BUS 2
DC bus 2
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Normal Operation
flight crew must rely on other sources of altitude information in the aircraft such as barometric altitude and glides lope.
The ERJ 170 Radar Altimeter System is a continuous service system, that provides the pilots with dependable, accurate altitude, up to +2,500 ft, above ground level (AGL) during the approach and landing phases of aircraft operation. In a single Radar Altimeter System installation the same radar altitude is displayed in both Primary Flight Displays (PFD`s) as a digital read out. In a dual Radar Altimeter System installation, the LH PFD displays the system #1 radar altitude and the RH PFD displays the system #2 radar altitude.
The pilots are able to select a Decision Height on the Minimums RA/BARO knobs located at the Guidance Panel. The DH is displayed on both PFD’s followed by a RA label. The DH indication is displayed on approach with a black box or a “MIN” annunciation as follows: a black box appears when the displayed radar altitude is less than or equal to DH plus 50 feet. A “MIN” annunciation appears when the displayed radar altitude reaches the DH.
Abnormal Operation With a single Radar Altimeter System installed upon failure of the system, a fail indication will be provided to the flight crew. With a dual Radar Altimeter System installation upon a single failure, only the affected side will lose the radar altitude read-out and a fail indication will be provided to the flight crew. In case of a main generator failure, the operational main generator will supply both DC Buses through the Transformer Rectifier Units (TRU) not affecting the radar altitude display. The APU generator can also take over the failed main generator and feed the DC Bus and the radar altitude will be not affected. In the case of a single TRU failure, the operational TRU will supply both DC Buses and also not affecting the radar altitude display. Only in the case when both TRU’s fail, the radar altitude will be not displayed, then the
last update: Nov06
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Chapter 34-31
Page 7
Figure 4: Operation
Display
Display
Controller
Controller
Decision Height selected on the Minimums RA/BARO knobs located on Guidance Panel
Current RA Altitude Pilot Entered DH
Decision Height (DH) range: 5 to 999 feet
last update: Nov06
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System and Component Function The ERJ 170 Radio Altimeter R/T transmits radio frequency pulses (4300 MHz) via the transmitting antenna, measures the time until the reflected signal is detected via the receiving antenna, and then determines the aircraft AGL altitude above the terrain. The R/T contains power supplies, radio frequency transmitting, and receiving circuitry, and also timing circuits to determine aircraft AGL altitude. The Radar Altimeter Receive and Transmit Antennas are designed to function between 4.2 and 4.4 GHz. In a single Radar Altimeter System installation the same radar altitude is displayed in both Primary Flight Displays (PFD`s) as a digital read-out. In a dual Radar Altimeter System installation, the LH PFD displays the system #1 radar altitude and the RH PFD displays the system #2 radar altitude.
last update: Nov06
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Chapter 34-31
Page 9
Figure 5: Functional Description
Radar Altimeter absolute altitude range: -20 to +2500 feet Decision height annunciation on PFD alerts the flight crew when the RA minimum altitude is reached. Below the RA minimum set value, MIN is displayed on the PFD. Preset altitude trip points alerts the flight crew to ground potential.
Parameter Options can be saved for future use. Once saved, parameter options can be loaded into the configuration module
If the zero point was offset, the radar altitude will read a negative value which corresponds to the amount of offset.
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34-32 Radio Navigation Introduction The VHF (Very High Frequency) NAV (Navigation) system uses airborne equipment and ground stations to supply data for in-flight navigation, approach/landing, and area guidance functions.
Figure 1: VHF NAV Overview
MCDU 1
MAU 1
MCDU 2
MAU 2
ASCB
VOR/LOC ANT
EDS
MRC 1
G/S ANT
MB ANT
VHF NAV 1 MODULE
VHF NAV 2 MODULE
MRC 2
MB ANT SPLITTER
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General Description The VHF NAV system is a precision ILS (Instrument Landing System) that supplies this data: • VOR (VHF Omnidirectional Range) in-flight navigation, terminal navigation, and area guidance. • LOC (Localizer)/G/S (Glideslope) approach and landing guidance. • MB (Marker Beacon) distance-to-runway threshold. The VHF NAV system includes these components: • VHF NAV Module • VOR/LOC (VOR Localizer) Antenna • MB Antenna • MB Antenna Splitter • G/S Antenna The aircraft has two VHF NAV systems.Each system has one VHF NAV module, which is an LRM (Line Replaceable Module).For the VHF NAV 1 system, the module is installed in the MRC (Modular Radio Cabinet) 1, in the forward avionics compartment.For the VHF NAV 2 system, the module is installed in the MRC 2, in the middle avionics compartment. The VHF NAV system can be tuned by the MCDU (Multifunction Control Display Unit) or by the CCD (Cursor Control Device) and PFD (Primary Flight Display).The MCDU is the primary controller and the CCD and PFD are the secondary controllers.
The DC (Direct Current) ESS (Essential) Bus 1 supplies the power to the VHF NAV 1 system.The DC Bus 2 supplies the power to the VHF NAV 2 system through the SPDA (Secondary Power Distribution Assembly) 2.
Figure 2: VOR/ LOC/ GS/ Marker Beacon Schematic
D
A
E
B B
F C
DU
J1
C
D
A CONTROL PEDESTAL
ARINC 429
DISPLAY UNITS
(SSM 31-62-80)
FWD AVIONICS COMPT
(SSM 31-61-80)
MAU 1 (SSM 31-41-80)
MAU 2 (SSM 31-41-80)
CONTROL PEDESTAL CONTROL I/O MODULE
CONTROL PEDESTAL
ARINC 429
CONTROL I/O MODULE
MCDU 2
MCDU 1 BACKPLANE BUS
ARINC 429
(SSM 34-61-80)
(SSM 34-61-80)
GENERIC I/O MODULE
GENERIC I/O MODULE
ARINC 429
COCKPIT
MAU 2 BACKPLANE BUS
MAU 1
F
E
MAIN INST PANEL
CCD' S FWD AVIONICS COMPT
N CO BEA R 39 RKE Hz . MA 75M 113 7LM INC . D SSP WI S. USA P/N .410 TEMCA. . S/N R SYS 91 SO WORTH 136 SENATS FSC CH
AL
J2
LH CBP
MIDDLE AVIONICS COMPT NIC
DC ESS BUS 1 5
ASCB
SPDA 2 (SSM 24-61-80)
NIC
COCKPIT
VOR/ILS 1
DC POWER - M3 DC BUS 2
DIGITAL AUDIO PANELS
DIGITAL AUDIO BUS 1
DIGITAL AUDIO BUS 2
(SSM 23-51-80)
(SDS 34-32) (MPP 34-32-05)
FWD AVIONICS COMPT MRC 1 (SSM 34-02-80)
SPLITTER NIM
VERT STAB VHF NAV 1 MODULE (SDS 34-32) (MPP 34-32-01)
D VOR 1 / 2 ANT (SDS 34-32) (MPP 34-32-03)
MIDDLE AVIONICS COMPT MRC 2 (SSM 34-02-80)
F
CENTER FUS I (DOWN)
B MB ANT (SDS 34-32) (MPP 34-32-04)
NIM
BULKHEAD (FWD) PRESS
C
VHF NAV 2 MODULE (SDS 34-32) (MPP 34-32-01)
G/S ANT (SDS 34-32) (MPP 34-32-06)
MRC 1 last update: Nov06
28 VDC
28 VDC
CENTER FUS I
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Introduction The ADF (Automatic Direction Finder) system uses airborne equipment and ground stations to supply the aircraft BRG (Bearing) data for in-flight navigation, terminal navigation, and area guidance functions.
ADF Components The ADF module is an LRM (Line Replaceable Module) installed in the MRC (Modular Radio Cabinet).Its rear connector connects to the MRC backplane which supplies the RCB (Radio Control Buses) interface to the NIM (Network Interface Module).The NIM supplies the interfaces to the ASCB (Avionics Standard-Communication Bus), ARINC (Aeronautical Radio Incorporated)-429 bus, digital audio bus, and other aircraft systems. The ADF module also has a backshell connector and a 50 ohm coaxial connector on its front panel that connects to the ADF antenna. The ADF module contains these circuits: • AM (Amplitude Modulation) receiver • BFO (Beat Frequency Oscillator) • +28 VDC (Volt Direct Current) power supply • Audio circuitry The ADF module supplies the relative BRG data as ARINC-407 synchro data, DC (Direct Current) sine/cosine data, and RS-422 data.
The ADF module has two bandwidths: • A narrow-band mode to reduce unwanted noise during navigation. • A wide-band mode to increase the quality of voice signals. NOTE: For increased BRG accuracy, the narrow band mode is recommanded. The aircraft has two ADF modules installed.The ADF 1 module is installed in the MRC 1, in the forward avionics compartment.The ADF 2 module is installed in the MRC 2, in the middle avionics compartment.
Figure 3: ADF Schematic A ADF 2 ANT ENNA
ADF 1 ANT ENNA
A
FWD AVIONICS COMPT
FWD AVIONICS COMPT
MAU 1 (SSM 31-41-80)
MAU 2 (SSM 31-41-80)
BACKPLANE BUS
BACKPLANE BUS
COCKPIT
MCDU 1
ARINC 429
MCDU 2
LH CBP
DC ESS BUS 5
ARINC 429
(SSM 34-61-80)
(SSM 34-61-80)
NIC
CONTROL PEDESTAL
CONTROL I/O MODULE
CONTROL I/O MODULE
NIC
CONTROL PEDESTAL
MIDDLE AVIONICS COMPT
ASCB
SPDA 2 (SSM 24-61-80)
1
ADF 1
DC POWER - M3 DC BUS
FWD AVIONICS COMPT CENTER FUS II (UPPER) ADF 1 ANTENNA (SDS 34-53) (MPP 34-53-02)
2
MIDDLE AVIONICS COMPT MRC 2 (SSM 34-02-80)
MRC 1 (SSM 34-02-80) DIGITAL AUDIO BUS 1
NIM
A
DIGITAL AUDIO BUS
2
NIM
CENTER FUS III (UPPER) 28 VDC
A
COCKPIT ADF 2 MODULE
ADF 1 MODULE (SDS 34-53) (MPP 34-53-01
ADF 2 ANTENNA (SDS 34-53) (MPP 34-53-02)
DIGITAL AUDIO PANELS )
(SDS 34-53) (MPP 34-53-01)
(SSM 23-51-80)
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Introduction The DME system calculates the time delay of radio pulses transmitted to and immediately received from a ground station.It uses the time data to calculate the distance from the ground station, ground speed, and time-to-station.The DME system also supplies the Morse code identification data. The DME is capable of tracking four channels to supply slant range, ground speed, time-to-station, and station identification.Two additional channels track station identification of preset channels for quick access.Frequency tuning is automatically paired with the VOR/LOC (VOR Localizer).The system operates over the frequency band of 962 to 1213 MHz. The DME interfaces with the EDS (Electronic Display System) and the FMS (Flight Management System) through the ASCB (Avionics Standard-Communication Bus).The DME identification audio is transmitted from the digital audio bus to the airborne audio system. The DME system includes these components: • DME module • DME Antenna
DME MODULE The DME module is one of the LRM (Line Replaceable Module)s installed in the MRC (Modular Radio Cabinet).The DME module connects to the NIM (Network Interface Module) through the RCB (Radio Control Buses) in the backplane of the MRC.The NIM supplies the interfaces to the ASCB and dig
ital audio bus.The DME module has a 50 ohm coaxial connector on its front panel for the antenna cable. The aircraft has two DME modules installed.DME 1 module is installed in MRC 1, in the forward avionics compartment.DME 2 module is installed in MRC 2, in the middle avionics compartment. The DME module contains these circuits: • • • • •
+28 VDC (Volt Direct Current) power supply Transmitter/receiver (transreceiver) Audio circuitry and IDENT outputs Mutual suppression circuitry Self-test circuitry
The DME module is a six-channel transreceiver.It monitors a maximum of four DME channels for distance, ground speed, and time-to-station, all at the same time.Two of the four DME channels are available for the FMS.The two remaining channels are available to control and show the distance, time-tostation, ground speed, and IDENT functions.The IDENT audio signals are transmitted on the digital audio bus to the DAP (Digital Audio Panel)s. The two preset IDENT channels have the decoded IDENT data immediately available.The VOR (VHF Omnidirectional Range) channel is a preset channel.After the VOR channel is set, the instant search function of the DME module immediately
34-41-01 EGPWS - WS location and modes Introduction The Enhanced Ground Proximity Warning System (EGPWS) uses aeroplane position, configuration and terrain database information to provide the flight crew with increased awareness of the terrain along the projected flight path. The system provides the flight crew with sufficient information and alerting to detect a potentially hazardous terrain situation and thus allow the flight crew to take effective action to prevent a controlled flight into terrain (CFIT) event.
Figure 1: Enhanced Ground Proximity Warning System (EGPWS)
The Enhanced Ground Proximity Warning System (EGPWS) uses: Position configuration Terrain database information
Goals: ncreased awareness of the terrain detect a potentially hazardous terrain situation prevent a controlled flight into terrain (CFIT) event
last update: Jun06
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EGPWS module The EGPWS is a module contained in the modular avionics unit (MAU) # 2 and it is powered via ESS DC BUS # 2.
Figure 2: EGPWS module
EGPWM
ESS BUS 2
MAU 2
last update: Jun06
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The EGPWS interfaces The EGPWS interfaces with the following systems and equipment: • Radio Altimeter - The radio altimeter provides altitude above the ground, how fast the altitude decreases due to the airplane sink or ground profile change and the validity signal. • ADCs - The ADCs provide uncorrected barometric altitude, corrected barometric altitude, computed airspeed, true airspeed, barometric altitude rate and static air temperature. • FMS - The FMS provides latitude, longitude, ground speed, true tracking, true heading and NAV mode. The same is applicable when the airplane is equipped with dual FMS. • Landing gear - The GPWS receives a discrete signal, which indicates gear down/locked condition. • Flap - The Flap Control Unit provides one discrete signal, which indicates whether or not flaps are in landing position. Aural Warning Unit - The aural warning unit receives the aural messages to be enunciated. It also provides a discrete signal to indicate that the glides lope advisory alert may be cancelled without any restriction.
Figure 3: The EGPWS interfaces
EMERG/ PRKG BRAKE
GND PROX TERR INHIB
LG WRN INHIB
GND PROX G/S INHIB
A
GND PROX FLAP OVRD
STEEP APPROACH
B
C
D
ARINC 429 CONTROL PEDESTAL
ARINC 429
MAIN INST PANEL
CCD 1
MAIN INST PANEL
DU 2 (MFD 1)
DU 1 (PFD 1)
(SSM 31-62-80)
MAIN INST PANEL
DU 4 (MFD 2)
(SSM 31-61-80)
(SSM 31-61-80)
CONTROL PEDESTAL
MAIN INST PANEL
CCD 2
DU 5 (PFD 2)
(SSM 31-61-80)
(SSM 31-62-80)
(SSM 31-61-80)
TERRAIN PICTURE BUS ASCB
MAU 1 (SSM 31-41-80)
BACKPLANE BUS
B
CONTROL PEDESTAL STEEP APPROACH
MAU 2 (SSM 31-41-80)
C
(SDS 34-41) (MPP 34-41-05)
ARINC 429
CONTROL PEDESTAL
GND PROX FLAP OVRD
D
GENERIC I/O
PROC 5
(SDS 34-41) (MPP 31-17-02) NIC
GENERIC I/O
ADA 2 FUNCTION (SSM 34-10-80)
GND PROX G/S INHIB
GPS 2 (SSM 34-56-80)
A
(SDS 34-41) (MPP 31-17-01)
PROC 4
NIC
FMS 1 FUNCTION (SSM 34-61-80)
NIC
PROC 3
MAIN INST PANEL
FMS 2 FUNCTION (SSM 34-61-80)
GND PROX TERR INHIB
EGPWS/WINDSHEAR MODULE (SDS 34-41) (MPP 34-41-01)
TERRAIN PICTURE BUS
ADA 1 FUNCTION
(SSM 34-26-80)
(SSM 34-10-80)
NIC
GENERIC I/O
GPS 1 (SSM 34-56-80)
IRU 1
MAIN INST PANEL
BACKPLANE BUS
PROC 1
FWD AVIONICS COMPT
MIDDLE AVIONICS COMPT
FWD AVIONICS COMPT
BACKPLANE BUS
ADA 3 FUNCTION (SSM 34-10-80)
FWD AVIONICS COMPT
MAU 3 (SSM 31-41-80)
(SDS 34-41) (MPP 34-41-03)
ARINC 429
ARINC 429 ARINC 429
ASCB CENTER FUS III
FWD AVIONICS COMPT
FWD AVIONICS COMPT
MIDDLE AVIONICS COMPT RAD ALT 2
MRC 1
ARINC 429
ARINC 429
(SSM 34-31-80) CENTER FUS III
IRU 2 MRC 2 (SSM 34-26-80)
(SSM 34-02-80)
(SSM 34-02-80) COCKPIT
RAD ALT 1 (SSM 34-31-80)
last update: Jun06
NIM
DIGITAL AUDIO BUS 1
DIGITAL AUDIO PANELS
DIGITAL AUDIO BUS 2
NIM
COCKPIT (SSM 23-51-80)
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34-42-01 WX RDR location and interface Introduction The airplane can be equipped with WU-660 or WU-880 Weather Radar System models. For additional information on functions and operations, refer to the manufacturer’s manual. The Weather Radar System is designed for detection and analysis of the weather during flight. The WU-660/880 Weather Radar is: • a lightweight, • X-band colour digital radar with display designed for weather location and analysis, and • for ground mapping.
Figure 1: The WU-660/880 Weather Radar
WU-660/880 Weather radar: Lightweight X-band color digital Ground mapping
last update: Nov06
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Ground Handling Heating and radiation effects of weather radar can be hazardous to life. Personnel should remain at a distance greater than the following to prevent exposure above safety levels: - 3.2 meters from the 18 inch radiating antenna. - 4.0 meters from the 24 inch radiating antenna (WU-880 only). Distances calculated using radiation exposure levels that equal or exceed 6 mW / cm 2. Honeywell recommends that operators follow the 6 mW / cm 2 standard per the IEEE Standard for Safety Level with Respect to Human Exposure to Radio Frequency Electromagnetic Fields 3 kHz to 300 GHz (IEEE C95.1 1991). The WX system is accessible through the nose of the aircraft.
Figure 2: Ground Handling
24 IN
last update: Nov06
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4 meters
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The Weather Radar System The Weather Radar System consists of an integrated receiver/transmitter/ antenna unit (RTA) and two virtual Weather Radar Controllers. The RTA is mounted in the nose of the airplane whereas the virtual controllers consists of the CCDs and the WX mode information displayed on the MFDs below the weather information. The Weather Radar System is powered by 28 VDC from SPDA 1, DC Bus 1. Should a power supply failure occur, the Weather Radar System will be inoperative, as there is no backup power source for this system.
Figure 3: Weather radar
Weather Radar Used for detecting weather up to 300 NM from the aircraft, ground mapping and detecting turbulence (WU-880 only). Displays storm intensity and position relative to the aircraft.
RTA
last update: Nov06
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Interfaces The RTA (WU -660/880) receives pitch and roll data for antenna stabilization from the Inertial System via ARINC 429 bus. The RTA (WU-880 only) receives Ground Speed and Inertial Altitude data from the Inertial Reference System via ARINC 429 bus to perform turbulence processing, and Altitude Compensated Tilt. The RAT accepts mode commands from the “virtual” WX Radar Controllers and the CCDs. Commands are delivered to the RTA on serial control buses via the MAU. The RAT outputs range, mode, gain and tilt information to the MAU on two Electronic Display System (EDS) control buses, and outputs scan converted data directly to Display Units (DU) 1,2,4, and 5, depending on display selections. The CCDs provide Range control to the MFDs via ARINC 429. The Weight-on-Wheels (WOW) discrete is provided by the Custom I/O in MAU 1 to the RTA to inhibit the transmitter on the ground. This is to prevent an X-band microwave radiation hazard. The WOW discrete must indicate an “in-air” condition before the transmitter is enabled. While inhibited on the ground, the WX mode is annunciated as Forced Stand-by (FSBY) in the WX mode field on the MFD. When WX or GMAP are selected while the aircraft is on the ground (WOW is true), the RTA mode will be annunciated as FSBY and the transmitter will be inhibited. FSBY Override (FSBY OVRD) enables the Weather Radar transmitter on ground, generally prior to take off, to monitor weather in the immediate area. FSBY OVRD is only available on ground: selection is inhibited in the air. If the WX Radar is active upon touchdown, it is forced into stand-by mode by the system. If the WX Radar is OFF upon touchdown, it remains in OFF and Forced Stand-by is not implemented. Manual override of Forced Stand-by is done selecting the FSBY OVRD check box in both virtual controllers.
The concentric knobs on the CCDs control tilt, range and gain. The outer knob controls range. The inner knob controls tilt when MFD is selected, except for when VAR WX Gain has been selected. When VAR WX Gain is selected, the inner knob controls gain. The CCDs communicate via ARINC 429 to the display units, which place the information on ASCB, to the Control I/ Os in MAU 1 and 2, and ultimately on RS-422 for use by the WX RTA. Range control by the CCDs is available full time. Tilt and gain commands are controlled through the CCD, displayed on the MFD and sent serially to the RTA. Tilt control is available whenever VAR Gain is not selected. The WX button on the Guidance/Control Display Panel (GP) allows the WX information to be displayed on the PFD. The GP communicates to the MAUs via RS 422 bus, and the PFD receives input from the GP through the MAUs.
Outputs The WX Radar pictures are supplied to the Display Units on the Left and Right WX Picture Bus. The Left Picture Bus is connected to the Pilot-side DUs (DU 1 and 2), and the Right Picture Bus is connected to the Copilot-side DUs (DU 4 and 5). The WX Picture Buses going to the Display units are serial busses that are isolated from the operating systems of the instruments. The RTA sends control words to the Control I/O modules in the MAUs via RS 422 to update information on the Display Units, including operating mode, selected range, RTA failure, and tilt angle.
Figure 4: Weather Radar Schematic
Honeywell Honeywell
DISPLAY
DISPLAY
CCD 1
CCD 2
A-429
PFD 1
A-429
MFD 1
MFD 2
PFD 2
IRS "Virtual" WX controller
"Virtual" WX controller
Left WX picture bus
RTA
Right WX picture bus
ASCB
NIC
I/0 WOW
MAU 1
last update: Nov06
RS-422
RS-422 Guidance Panel
NIC
NIC
MAU 3
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PFD WX Display WX Radar information will be displayed on the PFDs when WX is selected on the Guidance/Display Control Panel (GP). The display control portion of the GP on the pilot’s side will control the pilot’s PFD, and the copilot’s GP will control the copilot’s GP. The displays on the two PFDs are independent of each other.
Figure 5: PFD Weather Radar Display
WX Selectin
PFD
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Notes:
Figure 6: MFD WX Display Selection
MFD 1 Map
Pl an
MFD 2 Syst ems
Fuel
Map
Pl an
Syst ems
Fuel
Navaids
Airport AC
WPI Ident
AC
AC
AC
Progress Vert Prof TCAS
FUEL TEMP 32° C
DC
Weather
4000 LBS
TCAS
DC
Terrain
4000 LBS
WX
FUEL TEMP 32° C
4000
4000 LBS
LBS OFF
Honeywell
Honeywell
DISPLAY
DISPLAY
Checkl i st
TCAS
WX
Checkl i st
Enter keys
CCD 1
last update: Nov06
CCD 2
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Weather radar normal operation Display of weather data is available on the MFDs and PFDs. Selection of weather data display is accomplished via MFD using a soft key to select MAP menu. Selecting Weather from this menu will display the weather information and virtual controller on the MFD. When the Weather option is selected on the MAP menu following a cold start, the weather information will be displayed. Weather will initially display in a default OFF mode (OFF is the initial active mode selected on the virtual controller). After 45 seconds, when power up is completed, the pilot can select another available mode using the virtual controller. Functions are displayed in the virtual controller in white if available when a specific WX radar mode is selected, and displayed in gray if not available for the selected WX radar mode. Only one of the modes WX, GMAP, STBY and Off can be selected at one time. Availability of other functions will depend on which of these four modes is selected, and if the airplane is in the air or on the ground.
Figure 7: MFD Operation
Soft key
MFD
Not avalible (gray)
last update: Nov06
Avalible (white)
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WX on Ground When WX or GMAP are selected while the aircraft is on the ground (WOW is true), the RTA mode will be annunciated as FSBY and the transmitter will be inhibited. FSBY Override (FSBY OVRD) enables the Weather Radar transmitter on ground, generally prior to take off, to monitor weather in the immediate area. FSBY OVRD is only available on ground: selection is inhibited in the air.
Figure 8: WX On Ground
MFD 1
MFD 2
Forced Standby Override (FSBY OVRD) enables WX on the ground. Dual pilot action is required to override FSBY.
last update: Nov06
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A. Test Function (TEST) The test function is selected on the Multi-Function Control Display Unit (MCDU) avionics test page. the test is used to select a special test pattern to allow verification of system operation; 100-mile range is automatically selected. When the test is initiated, the TEST function is displayed on the MFD as a test pattern made up of Colour Bands. These bands are a series of green/yellow/red/magenta/white bands, which indicate that the signal to colour conversion circuits are operating normally. When TEST is selected in one of the MCDUs, both sides go to TEST and a green “TEST” is shown on the WEATHER annunciation field. The modes selected on the virtual controller just before TEST was initiated remain selected and the functions FSBY OVRD and SECT are available for selection. The WU-660 displays the TEST codes as text only and the WU-880 provides a test pattern and codes. TEST is displayed in the WX Mode Annunciation on the MFD adjacent to the navigation display. The colours on the test pattern during a right-to-left sweep are different than the colours on the test pattern during a left-to-right sweep.
Figure 9: Test Mode
The test function is selected on the MCDU avionics test page.
MCDU 1
MFD Map
Pl an
AC
Syst ems
Fuel
AC
Waiting for graphic FUEL TEMP 32° C
DC
4000 LBS
PERF
NAV
PREV
MENU
DLK
NEXT
FPL
PROG
DIR
TRS
RADIO
4000 LBS
BRT DIM
+-
TCAS
W X
Checkl i st
SP
The bands are a series of green/yellow/red/magenta/white bands, which indicate that the signal to color conversion circuits are operating normally
last update: Nov06
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RTA The RTA is an integrated unit that contains the receiver, transmitter, and antenna. The RTA is installed in the nose of the aircraft. An internal fan keeps the RTA in a safe temperature range. The RTA receives power from the aircraft’s 28 VDC (Volt Direct Current) bus through the SPDA (Secondary Power Distribution Assembly) 1. The flat plate antenna transmits and receives radar signals for the receiver transmitter. The integrated electronics contained in the base of the unit has the necessary circuitry to transmit, receive and process WX signals. The RTA does scan conversions of the radar signals and controls the interfaces to other system components, and the MFD and PFD. The RTA also includes the antenna positioning and control mechanism for the antenna flat plate. The RTA transmits and receives X-band radio frequency signals for the weather detection and GMAP functions. The RTA has the properties that follow: • Receiver The receiver is a low-noise, microprocessor-controlled circuit. The antenna receives radar echo signals and applies them directly to the receiver. The receiver has an adjustable video bandwidth that changes with the pulse width and range selections. The receiver also has manual and automatic gain controls. • Transmitter The transmitter operates at 9375 MHz (+/- 40 MHz). The transmitter sends radar signals directly to the antenna flat plate. The transmitter uses a magnetron circuit that supplies 10 kW of power to the antenna. The transmitter has an adjustable pulse width of 1.0 and 2.0 us that changes with the range and mode selections. • Antenna The antenna does a full symmetrical scan every 10 seconds in sectors of 60 degrees (sector scan sweep mode) or 120 degrees (normal sweep mode) in azimuth. The sector scan sweep has a range of 30 degrees for each side. Normal sweep has a range of 60 degrees for each side. The antenna is an electromechanically-controlled, movable flat plate with a 24 in diameter.
The RTA has two switches set in the base on each side of the electrical connector. The switches are labelled XMIT and SCAN, and are located on the left and right sides of the connector. The normal position for both switches is ON. The function of the switches are as follows: • XMIT Switch When the XMIT switch is in the OFF position, the transmitter will not operate. This is a safety feature. This is normally used for ground maintenance checks and operation of the radar. The OFF position stops the antenna from radiating microwave energy. • SCAN Switch When the SCAN switch is in the OFF position, the antenna will not scan. This is a safety feature. This is normally used for ground maintenance checks. The figure WEATHER RADAR SYSTEM - COMPONENT LOCATION provides further data on the preceding text.
Figure 10: RTA
last update: Nov06
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34-44 Lightning Sensor System Introduction LIGTHNING SENSOR SYSTEM The Lightning Sensor System detects and locates areas of lightning activity in a 200 NM radius around the airplane. The system displays lightning rate of occurrence and position relative to the airplane.
20
0
N
m
Figure 1: Lightning sensor system
last update: Jun06
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The Lightning Sensor Processor LOCATION The Lightning Sensor Processor is located in the forward electronics bay.
Figure 2: Lightning sensor processor location
The Lightning Sensor Processor
last update: Jun06
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LSS Interfaces Lightning sensor Antenna: • sends magnetic and electric analogue data associated with lightning activity. Modular Avionics Module (MAU) 1 and 2: • provides to the LSS information to be displayed on the MFDs. Modular Avionics Unit (MAU) 3: • provides present position which enables the LSS to accurately position the lightning rate symbols over the same geographic location regardless of how the aircraft is manoeuvring. Virtual Controller: • control information from the MFD is sent to MAUs 1 and 2 where it is converted to control discretes for application to the LSP. Transmit Inhibit Relay: disables the LSS receiver during HF transmissions. This prevents false lightning displays caused by the HF radio. The relay is enabled when HF is selected on an Audio Panel and the corresponding Push-To-Talk switch is pressed.
Figure 3: LSS Interfaces
Map
Pl an
Syst ems
Fuel
MFD 1 ASCB
HN LSS ANTENNA
HW
LSS A429 BUS LAN
F +12VDC TEST LOOP PILOT´S AUDIO PANEL
LIGHTNING
VHF1
VHF2
VHF3
HF
SAT
NAV1
NAV2
NAV3
ADF1
ADF2
PA
DME1
DME2
CABN
MKR
SELCAL
SPKR
BKUP VOL
EMER
ID
INPH
RAMP
HDPH
MIC
VOL
VHF1: 47 NORM BKUP
Honeywell
AUTO MASK
CO-PILOT´S AUDIO PANEL
SENSOR
VHF2
VHF3
HF
SAT
NAV1
NAV2
NAV3
ADF1
ADF2
ID
DME1
DME2
MKR
PA
SPKR
INPH
HDPH
VOL
SELCAL BKUP VOL
MIC
NORM BKUP
AUTO MASK
HF 2 PTT
PROCESSOR
WX
LSS A429 BUS RIGHT
MAU 2
TRANSMIT INHIBIT
Checkl i st
VIRTUAL CONTROLLER 1
CENTRAL DISCRETES Map
+28VDC
MIC
VHF1
HF 1 PTT
TRANSMIT INHIBIT RELAY
TRANSMIT INHIBIT RTN
TCAS
CENTRAL DISCRETES
MIC
VOL
MAU 1
CENTRAL PURPOSE BUS
Pl an
Syst ems
Fuel
MAU 3
MFD 2
EMER
CABN
RAMP
VOL
VHF1: 47
TCAS
WX
Checkl i st
Honeywell
VIRTUAL CONTROLLER 2
last update: Jun06
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34-52 Transponder system
The top and the bottom Transponder (XPDR) antennas are installed on the top and bottom of the forward fuselage.
General Description/Component Location The Transponder (XPDR) system is a dual (two-channel), Mode S diversity system. It transmits the identification code and barometric altitude of the aircraft. The aircraft and Air-Traffic Control-Radar Beacon-System (ATCRBS) ground stations use this data to prevent aircraft collisions. The Transponder (XPDR) system operates with the Traffic Alert and Collision Avoidance System (TCAS). The Transponder (XPDR) system includes these components: • Transponder (XPDR) Modules TRANSPONDER MODULE - Removal/Installation • Transponder (XPDR) Top and Bottom Antennas TRANSPONDER TOP/BOTTOM ANTENNAS -
Removal/Installation The aircraft has two Transponder (XPDR) modules and two pairs of top and bottom antennas. The Transponder (XPDR) 1 module is installed in the Modular Radio Cabinet (MRC) 1, and the Transponder (XPDR) 2 module is installed in the Modular Radio Cabinet (MRC) 2. The Modular Radio Cabinet (MRC) 1 is installed in the forward avionics compartment, and the Modular Radio Cabinet (MRC) 2 is installed in the middle avionics compartment.
The Transponder (XPDR) system uses airborne equipment and Air-Traffic Control-Radar Beacon-System (ATCRBS) ground stations. The airborne Transponder (XPDR) transmits at 1090 Megahertz (MHz). The Air-Traffic Control-Radar Beacon-System (ATCRBS) ground station transmits at 1030 Megahertz (MHz).
STANDBY The Transponder (XPDR) system is energized but does not transmit the altitude data. The stand-by mode is set on the RADIO page 1/2.
The Air-Traffic Control-Radar Beacon-System (ATCRBS) ground station transmits the radio pulses (interrogations) to the aircraft. When the aircraft receives the radio pulses, it immediately transmits the reply back to the AirTraffic Control-Radar Beacon-System (ATCRBS) ground station. The AirTraffic Control-Radar Beacon-System (ATCRBS) ground station then measures and uses the time delay result to calculate the distance to the aircraft (±500 feet). The angle at which the Air-Traffic Control-Radar Beacon-System (ATCRBS) antenna receives the reply gives the azimuth of the aircraft.
last update: Jun06
ALT-OFF The Transponder (XPDR) system transmits the reply in Mode A and Mode S, but it does not transmit the altitude data. ALT-ON The Transponder (XPDR) system transmits the reply in Mode A, Mode C, and Mode S, and transmits the altitude data. TA The Traffic Alert and Collision Avoidance System (TCAS) is in the Traffic Advisories (TA) mode. TA/RA The Traffic Alert and Collision Avoidance System (TCAS) is in the Traffic Advisories (TA)/Resolution Advisory (RA) mode.
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Chapter 34-52
Page 3
Figure 2: Transponder System Schematic XPDR 1 BOTTOM ANTENNA
B
XPDR 2 BOTTOM ANTENNA
A
XPDR 2 TOP ANTENNA
XPDR 1 TOP ANTENNA
A
FWD AVIONICS COMPT
FWD AVIONICS COMPT
MAU 1 (SSM 31-41-80)
MAU 2 (SSM 31-41-80)
BACKPLANE BUS
MCDU 1
MCDU 2
(SSM 34-61-80)
ARINC 429
(SSM 34-61-80)
NIC
ARINC 429
CONTROL PEDESTAL CONTROL I/O MODULE
CONTROL I/O MODULE
CONTROL PEDESTAL
NIC
B
BACKPLANE BUS
CENTER FUS I
CENTER FUS I XPDR 1 TOP ANTENNA (SDS 34-52) (MPP 34-52-02)
34-43 Traffic Alert and Collision Avoidance (TCAS) General The Traffic Alert and Avoidance System (TCAS) is an independent airborne System, which does not relay on ATC for control or coordination. It detects unsafe traffic conflicts with other transponder-equipped aircraft and assists the flight crew in avoiding intruders inside a protected airspace. This is accomplished by interrogating the Mode A, Mode C and Mode S transponders, tracking responses, and providing advisories to the flight crew of the vertical separation from intruders.
last update: Jun06
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Chapter 34-43
Page 1
Figure 1: The Traffic Alert and Avoidance System (TCAS)
last update: Jun06
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TCAS System The TCAS is an on board advisory system to act as a backup to the ATC radar and the “see and avoid” procedures. By computing the closure rate and altitude of all transponder-equipped aircraft in the surrounding airspace, TCAS can anticipate a potential midair collision much before this has a chance to materialize. TCAS continuously plots local air traffic on the associated displays and in the event of a conflicting path, gives to the pilot the correct avoidance manoeuvre by changing the vertical speed, that is, the TCAS computer calculation determines the manoeuvre that must be avoided and/or the climb or descent rate in order to escape from a potential collision. If the intruding aircraft is also equipped with TCAS, the two systems can communicate their mutual intentions through the Mode S transponders. The coordinated advisories that result allow the two pilots to execute complementary avoidance manoeuvres. Vertical guidance to avoid midair collisions is accomplished by interrogating the Mode A, Mode C and Mode S transponders of potential threat aircraft, tracking their responses, and providing advisories to the flight crew to assure vertical separation. Two levels of advisories are provided: • Traffic Advisories (TAs) which indicate the range, bearing and relative altitude of the intruder to aid in visual acquisition of the intruder. • Resolution Advisories (RAs) which indicate what vertical manoeuvre must be performed or avoid to assure safe separation. Each type of advisory is accompanied by a corresponding aural messages generated by the TCAS computer and sent to the MRC in order to be broadcast to the cockpit.
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Figure 2: TCAS System
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TCAS Computer The TCAS Computer is located in the forward avionics bay, which is pressurized, and temperature controlled. The TCAS Computer is mounted in an ARINC 600 4-Modular Component Unit (MCU) mounting tray.
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Figure 3: TCAS Computer Location
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TCAS Antenna The TCAS directional antenna is a four-element capable of transmitting in four selectable directions at 1030 Mega Hertz (MHz). The antenna is capable of receiving replies from all directions simultaneously with bearing information at 1090 MHz, using amplitude-ratio monopulse techniques. The TCAS omni-directional antenna is capable of transmitting and receiving frequencies from 960 to 1220 MHz. The TCAS Computer uses the directional antenna in order to display intruder bearing. The omni-directional antenna does not provide bearing information.
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Figure 4: TCAS Antenna Location
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Radar Altimeter Interface The TCAS computer accepts inputs from the Radar Altimeter System via ARINC 429 low speed input busses. If a second Radar Altimeter is installed, the TCAS computer will receive radar altitude from both Radar Altimeter Systems. The TCAS uses the Radar Altimeter #1 signal unless it goes invalid, in which case it will automatically start using the Radar Altimeter #2 signal.
MRC Interface The TCAS computer contains two analog audio outputs that provide TCAS aural traffic advisories and resolution advisories, to the MRC #1 and #2 in order to be broadcast to the cockpit.
Mode S Transponder Interface The TCAS computer contains a set of ARINC429 high-speed busses for communication with two mode S diversity transponders in the MRC#1 and #2. It uses ARINC 718/735 communications protocol (2 inputs, 2 outputs). The TCAS receives through the active Mode S Transponder mode control from the MCDUs and CCDs and also barometric altitude from the Air Data System (ADS).
ADMS Interface The TCAS computer contains a set of ARINC429 low speed busses for communication with the ADMS contained in the MAU, which allows fault and troubleshooting information for the system to be displayed and retrieved by maintenance personnel.
Display Interface Bus (MAUs and EDS) The TCAS computer has two sets of ARINC output busses for display of traffic and resolution advisories. The TA/RA Display #1 and #2 busses are high speed ARINC 429 busses that contain both traffic information and resolution advisory information For each bus, a valid discrete input is provided that indicates whether the display is functional.
last update: Jun06
These buses are connected to the Control Input/Output (I/O) 1 on MAU 1 and to Control I/O 2 on MAU 2 in order to be displayed on the displays (PFD and MFD).
DVDR Interface The TCAS RA are recorded on the Flight Data Recorder (FDR) portion of the DVDR as a mandatory parameter via the MAU’s interface. Also TCAS aural messages broadcast to the cockpit are recorded in the Cockpit Voice Recorder (CVR) portion of the DVDR via an analog path from the Audio Control Panels (ACP).
Discretes Interface · Weight on Wheels (WOW) Used for enable the TCAS computer for on board software loading and also to allow the system to go into the expanded test mode. This signal is also used to avoid grounded aircraft being displayed as an intruder. · Landing Gear Down Used for enable the TCAS computer for onboard software loading. · Warning Inhibits Used to revert automatically to TA ONLY when higher priority advisories occur (e. g.: Enhanced Ground Proximity Warning System (EGPWS))
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Figure 5: TCAS Schematic
MAIN INST PANEL
CONTROL PEDESTAL CCD' S
DISPLAY UNITS
(SSM 31-62-80)
(SSM 31-61-80)
ASCB FWD AVIONICS COMPT FWD AVIONICS COMPT
CONTROL PEDESTAL
CONTROL PEDESTAL
MAU 2 (SSM 31-41-80)
MAU 1 (SSM 31-41-80)
BACKPLANE BUS
MCDU 2
MCDU 1
BACKPLANE BUS
(SSM 34-61-80)
(SSM 34-61-80)
AURAL WARNING MUTE ARINC 429
ARINC 429
TCAS TOP DIR ANTENNA (SDS 34-43) (MPP 34-43-02)
NIC
CONTROL I/O MODULE
GENERIC I/O MODULE
CUSTOM I/O MODULE
CMC (SSM 45-45-80)
NIC
AURAL ADV
GENERIC I/O MODULE
ARINC 429
ARINC 429
CONTROL I/O MODULE
ASCB
FWD FUS (UP)
FWD AVIONICS COMPT L GENERAL PURPOSE LDG RETRACTED ARINC 429 FWD AVIONICS COMPT
SYNTH VOICE FWD AVIONICS COMPT SPDA 1 (SSM 24-61-80) DC POWER - M5 DC BUS 2
last update: Jun06
28 VDC
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TCAS Displays on Multi-Function Display (MFD) A TCAS Map Overlay and a TCAS Zoom format are available for display on the MFD The two formats are mutually exclusive. Both formats have the following attributes: • • • •
Controls the vertical sensitivity No bearing TAs and RAs Relative or Absolute altitude display Overlay of TCAS traffic advisories RA, TA, Proximate Traffic (PT) and Other Traffic (OT)
Pilot and co-pilot TCAS controls on the MFD’s are totally independent from each other. That means each pilot can control the TCAS as desired independent of the selected controls on the other side. The traffic symbols displayed are limited to the 8 highest priority intruders in order to avoid clogged displays with low priority intruders. This display limitation is performed via a pin strapping and is applicable to the Map Overlay and to the TCAS Zoom format. The TCAS Zoom format has a dedicated display with unique range control of the TCAS data. The TCAS map overlay uses overall range control. The TCAS Zoom Format is displayed when selected by activation of the TCAS menu button. Additionally, when a TA or RA is encountered and the Map Format is not in view at a range of less than 50 Nautical Miles (NM) with the TCAS Map Overlay displayed, the TCAS Zoom Format automatically pops into view (i.e., TCAS Auto Pop Up occurs). The TCAS Zoom Format has display priority over the Weather virtual controller or Checklist format when an Auto Pop Up occurs.
last update: Jun06
When the TCAS Zoom Format is activated, the TCAS Overlay on the Map Format (if selected for display) in the upper MFD Window, is removed from display. The TCAS Overlay on the Map Format remains removed until the TCAS Zoom Format is deactivated. Upon deactivation of the TCAS Zoom Format, the TCAS Map Overlay (if previously displayed) is reactivated. The TCAS Zoom Format is deactivated by the activation of Checklist format at any time by selecting the virtual button. A fixed Range Ring is displayed when the TCAS Zoom Format is displayed and provides a spatial reference for the distance of the displayed intruders. The Range Ring is positioned in the centre of the TCAS Zoom Format horizontally and vertically positioned so the top of the ring is in view at the top of the TCAS Zoom format, but only approximately 240 degrees of arc are shown. The Range Read-out is provided on the lower right side of the Range Ring and has a trailing NM label. The Range Read-out displays the TCAS Zoom Range selected by the Cursor Control Device (CCD) Inner Knob when the range function is enabled on the TCAS Virtual Controller. The CCD Inner Knob Icon is displayed to the left of the TCAS Zoom Range Read-out when the CCD Inner knob is capable of setting that range. Only one rate of adjustment exists for the TCAS Zoom Range and the value increments/decrements by “one Range increment” per knob click. Clockwise rotation increases the TCAS Zoom Range value, counter clockwise rotation decreases the value. The TCAS Zoom Range is a simple linear scale in which the maximum and minimum values represent the end of the scale. When the current TCAS Zoom Range value reaches the maximum or minimum, the read-out stays at that value if the knob continues to rotate in the same direction, with no further indication that the limit has been reached. Regardless of the number of knob clicks that occur at the end of the scale, which go “past the end of scale,” the first knob click in the opposite direction will begin increment/decrement (as appropriate for scale end) back through the available TCAS Zoom Ranges.
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Figure 6: TCAS MFD Selection
TCAS on Map menu selects TCAS Map Overlay for display on MFD.
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TCAS Displays Note: The TCAS Zoom Range selection affects the TCAS Zoom Format only. The Three NM Range Ring provides a unique symbolic reference to determine the proximity of the Traffic Targets. The Three NM Range Ring is a ring of twelve small circles (or dots) positioned in the centre of the TCAS Zoom Format, placed in a radius of three nautical miles around the Own Aircraft Symbol. The circles are arranged so that one circle is positioned every 30 degrees (0o, 30 o, 60 o, etc.) The circles placed at 0 o, 90o, 180 o and 270 o are larger in diameter than the remaining circles. The diameter of the Three NM Range Ring is scaled to reflect the selected TCAS Zoom Range. The Three NM Range ring is displayed on the TCAS Zoom Format only when the TCAS Zoom Range is 6NM, 12NM or 20NM. When displayed, the Three NM Range ring is always correctly spatially positioned with respect to the Aircraft symbol and the current TCAS Zoom Range, but the unique symbolic appearance of the Three NM ring precludes the need for a label. The range of the TCAS zoom format defaults to 6 NM each time the TCAS Zoom format is displayed in the Lower MFD Window. And is automatically reset to 6NM if the TCAS Zoom format is in view at a range greater than 12 NM when an RA or a TA occurs.
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Figure 7: TCAS MFD Controller
Deselect TCAS on Map menu to select Zoom for display on MFD.
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TCAS Display Cont. The TCAS Zoom Format is displayed when selected by activation of the TCAS menu button. Additionally, when a TA or RA is encountered and the Map Format is not in view at a range of less than 50 Nautical Miles (NM) with the TCAS Map Overlay displayed, the TCAS Zoom Format automatically pops into view (i.e., TCAS Auto Pop Up occurs). The TCAS Zoom Format has display priority over the Weather virtual controller or Checklist format when an Auto Pop Up occurs. When the TCAS Zoom Format is activated, the TCAS Overlay on the Map Format (if selected for display) in the upper MFD Window, is removed from display. The TCAS Overlay on the Map Format remains removed until the TCAS Zoom Format is deactivated. Upon deactivation of the TCAS Zoom Format, the TCAS Map Overlay (if previously displayed) is reactivated. The TCAS Zoom Format is deactivated by the activation of Checklist format at any time by selecting the virtual button. Upon on-ground power up, the TCAS Virtual Controller defaults to an Altitude Range Display of Normal, a Traffic Altitude Display of Relative and a Range Display of 6 NM. On in-air power up, the TCAS Virtual Controller reverts to the last selection for each.
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Figure 8: TCAS Controls
Range: selectable using CCD. ABS:
Changes intruder flight level - actual level or difference to own A/C.
Norm: +/- 2700. Expanded: +/- 12700.
TCAS CONTROL PANEL
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Figure 9: TCAS indications
RA (20 - 30 Secs)
TA Proximety Traffic +10
TA (35 - 45 Secs)
Other Traffic
Traffic Detected no Bearing Available last update: Jun06
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TCAS Displays on Primary Flight Display (PFD) A resolution advisory (RA) is an automatic display indication given to the pilot recommending a manoeuvre to increase vertical separation relative to an intruding aircraft.
the target zone location as follows: Red - avoidance zone; Green - fly to zone. When there is a RA condition, the flight director command bars shall be removed. The zones represent pitch-angles that should be left at once and pitch-angles to "fly to".
TCAS Resolution Advisory Commands are composed of either one or two Avoidance Zones and up to one Fly-To Zone. The commands provide Flight Path Vector (FPV) guidance information to the flight crew to recommend or prohibit a manoeuvre and prevent hazardous encounters with intruding aircraft. The Up Avoidance Zone, when displayed, extends from the top of the ADI to a FPV target based on current ground speed and vertical speed corrective guidance. The Down Avoidance Zone, when displayed, extends from the bottom of the ADI to a FPV target based on current, ground speed and vertical speed corrective guidance. The Up Avoidance Zone is displayed when a down advisory (descend corrective) is received. A Down Avoidance Zone is displayed when an up advisory (climb corrective) is received. When either single corrective is received, the Fly-To Zone is displayed on the end of the Avoidance Zone symbol unless a Preventive is indicated. If a Preventive command is indicated, the FlyTo Zone symbology is suppressed. When both a corrective and a preventive are received simultaneously, the Fly-To Zone is displayed between the Avoidance Zones. The Fly-To Zone symbology is compressed as the preventive and corrective commands begin to merge. The Fly-To Zone is compressed until a minimum Fly-To Zone height remains. The minimum Fly-To Zone height is compressed no further to ensure that a flyable command is displayed. To aid in the Pilot’s compliance with the corrective and/or preventive, coloration of the Flight Path Angle (FPA) Symbol, Flight Path Vector Speed Error Tape and Flight Path Angle Acceleration Pointer is provided. If the FPA is in the avoidance zone, the symbology will be coloured RED, otherwise will be coloured GREEN. The avoidance and fly to zones shall rotate with the pitch scale under roll conditions. The aircraft symbol colour shall be a function of
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Figure 10: PDF Resolution Advisory (RA)
PFD Resolutionary Advisory (RA) Avoidance Zones
last update: Jun06
Fly-to-Zone
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Self Test There are two ways to perform a self test on the system. The first one is via the MCDU and the second one is directly on the front panel of the computer. ·
– a) All TCAS computer unit front panel lamps come on for a 3 second lamp test. – b) The TCAS PASS green lamp then comes on for a 10 seconds display period, and then goes off. – c) No red fault lamps come on during the 10 seconds period.
Self test using the MCDU’s
Access the TCAS test page and activate the TCAS self-test function by pressing the associated LSK. The following events will occur: • a) Aural annunciation “TCAS TEST” is heard on the audio system. • b) A traffic test pattern consisting of four targets is displayed on the MFDs for eight seconds, consisting of the following: • An RA symbol at 3 o’clock, 2 NM, 200 ft above, level flight • A TA symbol at 9 o’clock, 2 NM, 200 ft below, climbing • A PT symbol at 3.6 NM, 33 degrees to right of the aircraft heading (approximately 1 o’clock), 1,000 feet below, descending • An OT symbol at 3.6 NM, 33 degrees to left of the aircraft heading (approximately 11 o’clock), 1,000 ft above, in level flight – c) Two RA test guidance commands indicating a Don’t Descend, Don’t Climb more than 2,000 fpm advisory are displayed as follows: • Red and green RA bands on both pilot & co-pilot VSIs • Red & green RA avoidance zones & fly-to zone on both ADIs – d) TCAS TEST annunciation is displayed on the PFDs and MFDs. – e) After eight seconds observe the following: • If the self test passes, “TCAS TEST PASS” is heard and the test patterns are removed (from ADIs, VSIs, and MFDs) • If the self-test fails, “TCAS TEST FAIL” is heard and TCAS FAIL is annunciated on the PFDs & MFDs • Self test using the TCAS computer front panel Activate the TCAS commanded self-test by pressing the PUSH TO TEST button on the RT-951 TCAS Computer front panel and verify:
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Figure 11: Self Test
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34-56 Global Positioning System (GPS) GPS System The GPS is an airborne, 12-channel, satellite receiver system that monitors the NAVSTAR satellite signals.It is a primary source of navigation data for the FMS. The GPS operates automatically and independently of other navigation systems.Its output data includes the aircraft position, velocity, time, and the satellite position, pseudo range, and delta range. The GPS has a RAIM (Receiver Autonomous Integrity Monitor) function that assure the accuracy of its output data.It also has a PRAIM (Predictive Receiver Autonomous Integrity Monitor) function that lets the FMS set the accuracy at specified locations and times for non-precision approaches. The interface between the GPS and the aircraft system is through the ASCB (Avionics Standard-Communication Bus), RS-422, and ARINC (Aeronautical Radio Incorporated)-429 bus. The GPS includes these components: • GPS Receiver Module • GPS Antenna The aircraft has two GPS receiver modules.One installed in the MAU (Modular Avionics Unit) 1 and the other installed in the MAU 3.There is one GPS antenna for each GPS receiver module.Both GPS antennas are installed on the top of the center fuselage II.
Figure 1: GPS System Schematic GPS 2 ANTENNA
A GPS 1 ANTENNA
A CENTER FUS II (UPPER)
FWD AVIONICS COMPT MAU 2 (SSM 31-41-80)
FWD AVIONICS COMPT
GPS 1 ANTENNA (SDS 34-56) (MPP 34-56-02)
BACKPLANE BUS
MAU 1 (SSM 31-41-80) BACKPLANE BUS CONTROL PEDESTAL
A
CONTROL PEDESTAL
GPS 1
ARINC 429
FMS 1 FUNCTION (SSM 34-61-80)
ARINC 429
(SSM 34-61-80)
NIC
(SSM 34-61-80)
CONTROL I/O MODULE
CONTROL I/O MODULE
NIC
GENERIC I/O MODULE
GPS 1 RECEIVER MODULE (SDS 34-56) (MPP 34-56-01)
ARINC 429
MCDU 2
EGPWM (SSM 34-41-80)
PROC 3 MCDU 1
MAU 1
CENTER FUS II (UPPER) GPS 2 ANTENNA (SDS 34-56) (MPP 34-56-02)
34-60 Flight management System (FMS) Introduction Flight Management System The PRIMUS EPIC® Flight Management System is an integrated system that provides data for the cockpit displays and Flight Control System. The system provides data for the cockpit displays and Flight Control System. The FMS serves as an aid to: • • • •
flight planning, navigation, performance, and database and redundancy management.
The Flight Management System provides complete flight planning capability including predictions of fuel and time. Once programmed, the FMS can provide control outputs to the autopilot system to fly the aircraft along the planned route, both laterally and vertically. The FMS also provides the EDS with the flight plan and status information for display. All data is transmitted to the FMS via ASCB-D. The FMS interfaces with the MCDUs and accepts input from the PFD and data loader.
Figure 1: The Primus Epic Flight Management System (FMS)
The FMS serves as an aid to: Flight planning Navigation Performance Database and redundancy management
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FMS Components The ERJ 170 FMS MCDUs are installed on the central console. There are two MCDUs, one unit located on each side of the centre console, next to each crew member seat, to the right of the pilot and the left of the copilot. Each MCDU is easily accessible and readable. The MCDUs are interchangeable. The Data Loader is located on the left side console, to the left of the pilot’s seat. The MAU is installed in the forward and central II electronic compartments. The components of the FMS receive power from the following sources: • The FMS 1 function resides on the NIC Processor Module in MAU 2, and receives power from DC bus 2. • The FMS 2 function (dual installation) resides on the NIC Processor Module in MAU 3, and receives power from ESS bus 1.
Figure 2: FMS 1 and 2
FMS 1
FMS 2
DC BUS 2
ESS BUS
MAU 2
last update: Nov06
MAU 3
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NIC Processor Module and Database Module The NIC Processor and Database Module are housed on MAUs 2 and 3.
Figure 3: FMS interfaces
MAU 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B
16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9
8 2 B 7 6 5 4 3 2 1 B
#
U S
C H
Power Supply 3 DC 1 AGM 1
MAU 2 C H
B
#
S
S
A 1 A 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1
CONTROL I/O 1 BRAKES (OUTBD) PSEM 1 AIOPA1
A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1
C H
8 7 2 B 6 5 2 B 4 3 2 1 B
#
C H
B
U S
C H
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DATABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2 Power Supply 1 DC 2
C H
B U S
S
S
A 1 A 1 A 1 A C H
16 1 B 15 14 13 12 1 B 11 10 1 B 9 1 B 1 B
8 1 B 7 6 5 4 3 2 1 B 1
B U S
PROC 1 = ADA 1, MW 1, UTIL 1, CAL/MCDU 1, CMS 1
B
#
U S
C H
Power Supply 2 DC 2 ENGINE VIBE GPS 2 PSEM 2
C H
B U S
A 2 2 A
FCM 3 A 2 GENERIC I/O 3 A 2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 NIC 5 (A) (ID = 33) CUSTOM I/O 2
Power Supply 2 ESS 2/DC 2 BRAKES (INBD) CONTROL I/O 2 AIOPA2
B
U
16 15 14 13 12 11 10 9
MAU 3
DATABASE
FMS 2
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FMS Schematic, Interfaces The FMS interfaces with the following systems/components: • GPS The FMS uses inputs from the GPS to calculate aircraft position and perform navigation functions. • IRS THE FMS uses inputs from the IRS to obtain aircraft position and perform navigation functions. Automatic Direction Finder (ADF) • The FMS interfaces with the ADF via the Radio System bus. The FMS provides a means for back-up tuning of the ADF. • CMF The FMS interfaces with the CMF via the ASCB.
Figure 4: FMS Schematic, Interface RADIO
1/2 COM2
SQ
COM1
123.2OO
123.2OO 118.6OO
118.6OO FMS
NAV1
114.8 DME H PXR 115.6
NAV2
117.4
AUTO
CCD' S
IDENT
TA/RA
STBY
CONTROL PEDESTAL
116.8 H123 XPDR 1471
TCAS/XPDR
C
D
E
COCKPIT
A
B
G
H
I
J
K
L
1
2
3
+/ -
M
N
O
P
Q
R
F
4
5
6
/
S
T
U
V
W
7
8
9
X
Y
Z
FWD AVIONICS COMPT
LH CBP
SPDA 1 (SSM 24-61-80)
DC BUS 1
0
MCDU 1
MAIN INST PANEL
ARINC 429 MOD. 18
5
ARINC 429
(SSM 31-62-80)
MIDDLE AVIONICS COMPT
COCKPIT
SPDA 2 (SSM 24-61-80)
LH CBP
DC ESS BUS 2
ARINC 429 MOD. 17
MCDU 2
5
DISPLAY UNITS
A
(SSM 31-61-80)
ASCB LAN ARINC 429
ARINC 429
ARINC 429
ARINC 429
28 VDC
fms 1
28 VDC
FWD AVIONICS COMPT CONTROL PEDESTAL
MCDU 1
A
MCDU 2
(SSM 34-61) (MPP 34-61-02)
MAU 2 (SSM 31-41-80)
A
BACKPLANE BUS
(SSM 34-61) (MPP 34-61-02)
ADA 2 FUNCTION
(SSM 34-10-80)
PROC 4
NIC
FMS 1 FUNCTION
NIC
(SDS 34-61) (MPP 31-41-04)
CONTROL PEDESTAL
EGPWM (SSM 34-41-80)
ARINC 429 ARINC 429
GENERIC I/O MODULE
CONTROL I/O MODULE
ARINC 429
CONTROL I/O MODULE
GENERIC I/O MODULE
GPS 1 (SSM 34-56-80)
CMC (SSM 45-45-80)
MW 1 FUNCTION
PROC 3 (SSM 31-53-80)
ADA 1 FUNCTION
(SSM 34-10-80)
NIC
PROC 1
MW 2 FUNCTION
BACKPLANE BUS
CONTROL PEDESTAL
(SSM 31-53-80)
MAU 1 (SSM 31-41-80)
FWD AVIONICS COMPT
PRINTER (SSM 23-24-80)
ARINC 429
FWD AVIONICS COMPT
(SSM 34-26-80)
ARINC 429
ARINC 429
LAN PORTS
ARINC 429 ARINC 429
ARINC 429
LAN
LAN
FWD AVIONICS COMPT
ARINC 429 ASCB FWD AVIONICS COMPT
MAU 2
ASCB
RS 422 IRU 1
ARINC 429 ARINC 429 RS 422
ARINC 429
MIDDLE AVIONICS COMPT
MIDDLE AVIONICS COMPT
MRC 1
T
IRU 2 (SSM 34-26-80)
MRC 2
(SSM 34-02-80)
BACKPLANE BUS MAU 3 (SSM 31-41-80)
GPS 2 (SSM 34-56-80)
PROC 5
GENERIC I/O MODULE
ADA 3 FUNCTION
(SSM 34-10-80)
(MPP 31-41-04)
FMS 2 FUNCTION
NIM
NIC
NIM
(SDS 34-61)
(SSM 34-02-80)
MAU 3
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Figure 5: Operation
Vert Prof on Map or Profile menus selects Vertical Profile flight path data for display
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Figure 6: Functional Description
FMS pages: Direct – To (DIR) inserts the following prompts on the ACTIVE FLT PLAN pages: Direct
CB
Pattern
DLK
TRS
ntercept
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Database The FMS has three data bases in its memory. There is a navigation database, a custom database, and an aircraft database. For the FMS to operate efficiently, these data bases, or portions thereof, must be accessible through direct memory addressing from the FMS Processor module. For this purpose, the FMS Processor module provides at least 8 Megabytes of nonvolatile Random Access Memory (RAM) dedicated to database storage to store the portions of the FMS data bases that require direct memory addressing. The Database module houses a copy of all three data bases.
• • • • • •
Navigation Database
•
This database contains data on NAVAIDS, airports and airways. The information in the navigation database is updated every 28 days. The database contains two consecutive effectivity cycles, and the correct database may be automatically or manually selected. Automatic selection occurs upon entry of a valid date. Dates are valid if they are greater than or equal to the expiration date of the current cycle.
•
The PRIMUS EPIC® Database module provides the storage capacity for at least 32 Megabytes of data for the FMS Navigation Database. There are three ways that data can be retrieved from the database. Ident searches are used when an operator requests a particular data item. Position searches are used to create lists of items within a specified proximity to a data item. NAVAID frequency searches are used to identify which station is being tuned by a radio. The navigation database contains the following types of data: • • • •
NAVAIDS -- worldwide high altitude VHF Navigation Aids DME only non-collocated VOR/DME TACAN only
• • •
non-collocated TACAN VORTAC collocated VOR/DME VOR only ILS/MLS -- worldwide Instrument Landing System (ILS) and Microwave Landing System (MLS). Airports -- worldwide airport geographic points for airports with International Civil Aviation Organization (ICAO) identifiers and one hard surface runway at least 4000 feet long. Airport Runways -- worldwide airport runways which are at least 4000 feet in length. Airport Procedures -- worldwide SIDs, STARs, and approach procedures. Named Way points -- worldwide named way points, intersections, and non-directional beacons (NDBs). Unnamed Way points -- way points computed in the process of path/terminator conversions. Airways -- worldwide high and low altitude airways.
Figure 7: Data Bases
FMS Databases: Navigation: contains data on NAVAIDs, airports, and airways. It is updated every 28 days.
Custom: contains flight plan and waypoint information entered by the flight crew. This database is not updated on a scheduled basis.
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Custom Database The custom database contains information entered by the pilot. This is where the pilot can create and store flight plans and way points. This database is not updated on a scheduled basis. The PRIMUS EPIC® Database module provides the storage capacity for at least 128 Kilobytes of data for the FMS Custom Database.
Aircraft Database The aircraft database contains all aircraft-specific performance parameters. After each flight, all learned data is saved to the file automatically. Thus, all performance data (learned and fixed) is contained in the file. The aircraft database file is only used for predictions in the FULL PERF mode. The aircraft database file also contains all aircraft- specific data needed for speed selection logic and MCDU displays. The PRIMUS EPIC® Database module provides the storage capacity for at least 32 Kilobytes of data for the FMS Aircraft Database.
Figure 8: Data bases (continued)
FMS Databases: Aircraft: contains all aircraft-specific performance paramters. After each flight, all learned data is saved to thisfile automatically. CB DLK
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Data Loading The data loader provides a means to load database information (navigation, aircraft, or custom) or save stored information Data bases can also be crossloaded between FMSs.
Abnormal Operation In the unlikely event that the FMS becomes inoperative in a dual FMS installation, the crew may rely on the remaining FMS for all of the functions performed by the FMS. If both FMS fail, or in the case of a single FMS installation, the crew may rely on other navigation instruments for the information normally provided by the FMS. In the event an FMS becomes inoperative and cannot be used to perform the required operations, a Crew Alerting Message (CAS) is posted on the Engine Indicating and Crew Alerting System (EICAS) display. For a dual FMS configuration, if one FMS fails, only the remaining FMS will be controlled by its corresponding MCDU.
Figure 9: DMU
Data Man agem ent Unit (DMU)
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34-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ 23-00 Standby ¦ Magnetic Compass ¦ System ¦
B ¦ 1 ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
B ¦ 1 ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Any combination of two IRS ¦ stabilized Compass Systems ¦ operate normally, and ¦ b) Airplane is operated with ¦ Dual Independent Navigation ¦ Capability and under ¦ Positive Radar Control by ¦ ATC on the enroute portion ¦ of the flight.
¦ 25-00 Head Up Guidance ¦ *** System (HGS)
D ¦ 2 ¦
¦ 1 ¦
¦ ¦
| ¦ | ¦
¦ ¦ ¦
D ¦ ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ approach minimums or operating ¦ procedures do not require its use.
| ¦ | ¦ | ¦
¦ ¦
¦ ¦
¦ ¦
¦ NOTE: ¦
| ¦ | ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 34-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 34 NAVIGATION ¦ ¦ ¦
¦ 11-01 Integrated ¦ Electronic ¦ Standby System
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ Deleted, Rev. 3. ¦ ¦
¦ 15-05 Total Air C ¦ 4 ¦ Temperature (TAT) ¦ ¦ Indications ¦
¦ 1 ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
C ¦ 4 ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided at ¦ least one SAT indication is ¦ available.
¦ ¦ ¦
¦ 15-07 Static Air C ¦ 4 ¦ Temperature (SAT) ¦ ¦ Indications ¦
MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 34-3 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 34 NAVIGATION ¦ ¦ ¦
¦ 31-00 Radio Altimeter ¦ System ¦ ¦ ¦ ¦ ¦ ¦
1) ERJ-170 Airplanes equipped with FADEC 4.12
¦ ¦ ¦ ¦ ¦ ¦ ¦
2) ERJ-190 Airplanes and ERJ-170 Airplanes equipped with FADEC 5.1 and on
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
¦ ¦
| ¦ ¦
C ¦ 2 ¦ ¦ ¦ ¦ ¦
¦ 1 ¦ ¦ ¦ ¦ ¦
¦ (M)Radio Altimeter 2 may be ¦ inoperative provided: ¦ a) System is deactivated, and ¦ b) Approach minimums or ¦ operating procedures do not ¦ require its use.
| | | | | |
¦ ¦ ¦ ¦ ¦ ¦
C ¦ 2 ¦ ¦ ¦ ¦ ¦ ¦
¦ 1 ¦ ¦ ¦ ¦ ¦ ¦
¦ (M)May be inoperative provided: ¦ a) System is deactivated, and ¦ b) Approach minimums or ¦ operating procedures do not ¦ require its use. ¦ ¦
| | | | | | |
¦ ¦ ¦ ¦ ¦ ¦ ¦
A ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (M)May be inoperative provided: ¦ a) System is deactivated, ¦ b) Approach minimums or ¦ operating procedures do not ¦ require its use, ¦ c) Ground Proximity Warning ¦ System (GPWS) Modes 1-4, ¦ Mode 5, Advisory Callouts ¦ and Windshear Mode are ¦ considered inoperative, ¦ d) Traffic Alert and Collision ¦ Avoidance System (TCAS) is ¦ considered inoperative, and ¦ e) Repairs are made within two ¦ flight days.
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 34-5 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 34 NAVIGATION ¦ ¦ ¦
¦ 41-00 Terrain Awareness ¦ and Warning ¦ System (TAWS) ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
| ¦ | ¦ | ¦
¦ 0 ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Alternate procedures are ¦ established and used, and ¦ b) Repairs are made within two ¦ flight days.
| ¦ ¦ ¦ ¦ ¦
A ¦ 4 ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Alternate procedures are ¦ established and used, and ¦ b) Repairs are made within two ¦ flight days.
| ¦ ¦ ¦ ¦ ¦
A ¦ 1 ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ May be inoperative provided: ¦ a) GPWS is considered ¦ inoperative, and ¦ b) Repairs are made within two ¦ flight days.
| ¦ ¦ ¦ ¦ ¦
1) Ground A ¦ 1 Proximity ¦ Warning System ¦ (GPWS) ¦ ¦
¦ ¦ ¦ ¦ ¦
a) Modes 1-4
¦ ¦ ¦ ¦ ¦
b) Test Mode
¦ ¦ ¦ ¦
c) Glideslope Deviation Mode(s) (Mode 5)
C ¦ ¦ B ¦ ¦
¦ 1 ¦ ¦ 0 ¦
¦ ¦ ¦ ¦
¦ ¦ | ¦ ¦
¦ ¦ ¦
d) Advisory Callouts
B ¦ ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
| ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Advisory callout not ¦ required by FAR, and ¦ b) Alternate procedures are ¦ established and used.
| | | | |
¦
¦
¦
¦ (Continued)
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 41-00 Terrain Awareness ¦ and Warning ¦ System (TAWS) ¦ (Cont'd)
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
| | | |
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
| | | | |
¦ ¦ ¦ ¦ ¦
C ¦ 1 ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Alternate procedures are ¦ established and used, and ¦ b) Windshear Detection and ¦ Avoidance System ¦ (Predictive) operates ¦ normally.
| ¦ ¦ ¦ ¦ ¦ | ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
C ¦ 1 ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Alternate procedures are ¦ established and used, and ¦ b) Takeoffs and landings are ¦ not conducted in known or ¦ forecast windshear ¦ conditions.
MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 43-00 Traffic Alert and B ¦ ¦ Collision ¦ ¦ Avoidance ¦ ¦ System II ¦ ¦ (TCAS II) ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦
¦ (M)May be inoperative provided: ¦ a) System is deactivated and ¦ secured, and ¦ b) Enroute or approach ¦ procedures do not require ¦ its use.
¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ (M)(O)May be inoperative provided: ¦ a) Not required by FAR, ¦ b) System is deactivated and ¦ secured, and ¦ c) Enroute or approach ¦ procedures do not require ¦ its use.
¦ ¦ ¦ ¦ ¦ ¦ ¦
C ¦ 2 ¦ ¦ ¦
¦ 1 ¦ ¦ ¦
¦ May be inoperative on non-flying ¦ pilot side. ¦ ¦
¦ ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Traffic Alert (TA) visual ¦ display and audio functions ¦ are operative, ¦ b) TA only mode is selected by ¦ the crew, and ¦ c) Enroute or approach ¦ procedures do not require ¦ its use.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) RA visual display and audio ¦ functions are operative, and ¦ b) Enroute or approach ¦ procedures do not require ¦ its use.
¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 34-7 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 34 NAVIGATION ¦ ¦ ¦
¦ 41-00 Terrain Awareness ¦ and Warning ¦ System (TAWS) ¦ (Cont'd)
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
| | | |
¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦ ¦ ¦ ¦ ¦ ¦ ¦
| | | | | | | | | |
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
C ¦ ¦ B ¦ -
¦ 1 ¦ ¦ 0
¦ ¦ ¦
| ¦ | ¦ | ¦
C ¦ 1 ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ ¦ ¦ ¦
| | | |
¦ 42-00 Weather Radar ¦ System
C ¦ ¦
¦ ¦
¦ Any in excess of those required by ¦ FAR may be inoperative.
¦ ¦
¦ ¦ ¦ ¦ ¦
B ¦ 1 ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ (M)May be inoperative provided: ¦ a) Antenna sweep is parallel to ¦ aircraft pitch axis, and ¦ b) Antenna tilt operates ¦ normally.
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ *** ¦ ¦ ¦
¦ ¦ ¦ ¦
2) Terrain System B ¦ 1 - Forward ¦ Looking ¦ Terrain ¦ Avoidance ¦ (FLTA) and ¦ Premature ¦ Descent Alert ¦ (PDA) ¦ Functions ¦ a) Terrain Display Functions 3) Runway Awareness & Advisory System (RAAS)
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 60-00 Flight Management C ¦ ¦ System ¦ ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
¦ ¦
D ¦ 2 ¦
¦ 1 ¦
¦ One may be inoperative provided ¦ procedures do not require its use.
¦ ¦
C ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be out of currency provided: ¦ a) Current Aeronautical Charts ¦ are used to verify ¦ Navigation Fixes prior to ¦ dispatch, ¦ b) Procedures are established ¦ and used to verify status ¦ and suitability of ¦ Navigation Facilities used ¦ to define route of flight, ¦ and ¦ c) Approach Navigation Radios ¦ are manually tuned and ¦ identified.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 34-9 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 34 NAVIGATION ¦ ¦ ¦
¦ 44-00 Lightning Sensor ¦ *** System
D ¦ ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ 51-00 Distance ¦ Measuring ¦ Equipment (DME) ¦ Systems
D ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ Any in excess of those required by ¦ FAR may be inoperative. ¦ ¦
¦ ¦ ¦ ¦
¦ 52-00 ATC Transponders B ¦ ¦ and Automatic ¦ ¦ Altitude ¦ ¦ Reporting Systems ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦
¦ May be inoperative provided: ¦ a) Enroute operations do not ¦ require its use, and ¦ b) Prior to flight, approval is ¦ obtained from ATC facilities ¦ having jurisdiction over the ¦ planned route of flight.
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
D ¦ ¦
¦ 1 ¦
¦ Any in excess of those required by ¦ FAR may be inoperative.
¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
1) Elementary A ¦ and Enhanced ¦ Downlink ¦ Aircraft ¦ Reportable ¦ Parameters not ¦ Required by ¦ FAR ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ May be inoperative provided: ¦ a) Enroute operations do not ¦ require its use, and ¦ b) Repairs are made prior to ¦ the completion of the next ¦ heavy maintenance visit. ¦ ¦
| | | | | | | |
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 53-00 Automatic ¦ Direction Finder ¦ (ADF) System
D ¦ ¦ ¦
¦ ¦ ¦
¦ Any in excess of those required by ¦ FAR may be inoperative. ¦
¦ ¦ ¦
¦ 56-00 Global ¦ Positioning ¦ System
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ procedures do not require its use. ¦
A and lane B both must be operational for the servos in that channel to be active and engaged.
General Description
AUTOPILOT
• • • • •
The FGCS comprises: Autopilot Yaw damper/turn coordination Automatic pitch trim FD
The FGCS is designed and built in dual channel architecture, configured in a master / slave arrangement. The master (or active) channel provides the computed control and monitoring functions. The slave (or stand-by) channel operates in a back-up mode as a ″hot″ spare. In the event that the active channel detects a failure, or it is otherwise disabled, the channel priority switches to the stand-by channel. The new active channel continues to provide the required functions, with no interruption to AFCS functionality during channel transition. When the original channel recovers from the invalid condition, it resumes the role of the stand-by channel. Assignment of priority alternates between the channels on power-up. Alternating channel priority is required to limit the exposure of either channel to a latent fault. The master AFCS/FGCS channel could be manually selected by using the MCDU (Multi function Control Display Unit) SETUP Page. Each AFCS module has failpassive/fail-safe processing, with no reliance on the other AFCS channel. No single fault within the AFCS can cause a hazard to the aircraft.
The autopilot supplies automatic pitch and roll guidance by sending commands to the elevator and aileron servos, through the CAN (Controller Area Network) bus. The commands start to move the cockpit control columns and control yoke in the commanded direction. Then the primary flight control system senses the cockpit control column/wheel position changes and directly commands the elevator and aileron control surfaces to carry out the autopilot pitch and roll guidance. The autopilot control authority is limited to the safe range of motions for the aircraft.
YAW DAMPER/TURN COORDINATION The yaw damper system provides rudder commands to improve damping of the dutch roll mode and to provide turn coordination. The AFCS supplies the rudder turn coordination to eliminate the excessive centripetal or centrifugal forces during a banked turn. The AFCS sends rudder turn coordination commands through the ASCB to the FCM (Flight Control Module)s. The FCM sends rudder turn coordination data to the ACE (Actuator Control Electronics) on the CAN bus. The AFCS processor receives from the FCM a verification of its command through the ASCB. The loss of the rudder turn coordination function does not cause the yaw damper to disengage.
The FGCS functions (autopilot, yaw damper/turn coordination, automatic pitch trim and flight director) are contained in the four AIOP (Actuator InputOutput Processor) modules. These four LRM (Line Replaceable Module) are installed in the MAU (Modular Avionics Unit) modules, and they are interfaced with the other MAU modules, using the ASCB (Avionics StandardCommunication Bus) data bus. The AIOP modules are also interfaced to the other, external to MAU, avionics and flight controls equipment and systems, as appropriate and required for the AIOP’s function. Two AIOPs operate in each channel. These modules are identified as lane A and lane B. The modules in these two lanes have the same software, but do separate, complementary, and similar functions that depend on the lane. The AIOPs in lane
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 1
Figure 1: AFCS Architecture
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
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170/190 MAINTENANCE TRAINING MANUAL
Automatic Pitch Trim The automatic pitch trim function positions the horizontal stabilizer surface to reduce the aerodynamic force held by the elevator. The automatic pitch trim function operates when the autopilot is engaged. The FGCS sends commands through the ASCB to the FCM. The FCM, through the CAN bus, sends commands to the HS-ACE (Horizontal-Stabilizer Actuator-Control Electronics). The HS-ACE sends commands to the HSA (Horizontal-Stabilizer Actuator). The HSA operates the horizontal stabilizer to supply the automatic pitch trim function. The FGCS uses both surface rate feedback and stabilizer positions in the automatic trim calculations.
Flight Director The FGCS calculates the FD guidance commands for display on the PFD (Primary Flight Display)s. The FD is selected using the SRC push button on the GP (Guidance Panel). Only one vertical mode and one lateral mode are allowed to be armed at one time.
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 3
Figure 2: Flight Guidance and Control System
FMS
AIR DATA SYSTEM
IRS
RADAR ALTIMETER
FGCS FUNCTIONS YAW DAMPER/ TURN COORDINATION
FLIGHT DIRECTOR
AUTOMATIC PILOT
VHF NAV
AUTOMATIC PITCH TRIM MACH TRIM
AIOP
AIOP
CAN BUS
AUTOPILOT SERVOS
ASCB
PFD
last update: Jun06
FLIGHT CONTROL SYSTEM
CAN BUS
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Chapter 22-10
Page 4
170/190 MAINTENANCE TRAINING MANUAL
Notes:
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 5
Figure 3: Flight Guidance and Control System
CONTROL YOKE
CONTROL YOKE
AILERON
ELEVATOR
RUDDER PEDALS ELEVATOR SERVO
LVDT
MAU'S
LVDT
AIOPS
FCMS
P-ACES
AILERON SERVO
ACTUATOR
ACTUATOR
ACTUATOR
ACTUATOR HYDRAULIC
CAN BUS
MECHANICAL
ELECTRICAL
last update: Jun06
ACTUATOR
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Chapter 22-10
Page 6
170/190 MAINTENANCE TRAINING MANUAL
Notes:
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FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 7
Figure 4: Flight Guidance And Copntrol System FGCS - Block Diagram
AIOP Modules (FGCS Function) The four AIOPs are installed in the MAUs. The MAU 1 contains the AIOP 1A and 1B. The MAU 2 contains the AIOP 2A. The MAU 3 contains the AIOP 2B. The AIOP contains these systems installed: • FGCS • TMS (Thrust Management System) • SWPS (Stall Warning and Protection System) The AIOPs do all of the necessary computations and data processing for the autopilot and yaw damper functions. The AIOPs collect the necessary data from the ASCB and from other FGCS data inputs. The AIOPs send position data to the servos through a bidirectional CAN data bus. The AIOP is a dualslot, single-lane module that contains two interconnected circuit cards. One is an I/O (Input/Output) card and the other is a processor card. The I/O card does the actuator control functions, sends and receives signals for the excitation/demodulation of the position sensors, and controls the functional enable/disable control lines. The processor card does the core processing functions for the FGCS and controls the ASCB interface. The AIOP uses the datas that follows: For IRS (Inertial Reference System): • Pitch angle • Roll angle • Pitch body rate • Roll body rate • Yaw body rate • Normal acceleration • Lateral acceleration • Longitudinal acceleration • Ground speed • True track angle • Flight path angle • Magnetic heading • True heading • Inertial vertical speed
last update: Jun06
For ADS (Air Data System): • Baro-corrected altitude • Altitude rate • Pressure altitude • Vertical speed • Mach number • CAS (Calibrated Airspeed) • TAS (True Airspeed) • Dynamic pressure • VMO (Maximum Operating Velocity) • Mmo (Maximum Mach Operation) For the other sensors and systems: • Monitor and warning system • Servos and position sensors • Integrated pitot static AOA (Angle of Attack) sensor (AOA data) • Proximity sensing system • APM (Aircraft Personality Module) • Flaps system data • Radio altitude • Radio navigation data • FMS (Flight Management System) The FGCS includes the components that follow: • Autopilot aileron servo • Autopilot aileron cable • Autopilot elevator servo • Autopilot elevator cable • AP/TRIM DISC push button • AP/FD TCS push button
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Figure 1: AFCS MAU Architecture
MAU 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B
16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9
8 7 6 5 4 3 2 1
2 B
C H
B U
#
S
CMC GPS 1 Power Supply 2 ESS 1 FCM 1 A 1
16 15 14 13 12 11 10 9
AIOPB1 PROC 1 NIC 1 (A) (ID = 1) FCM 2
AIOPA1
A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1
2 B
SPARE SPARE GENERIC I/O 2
U
C H
8 7 6 5 4 3 2 1
2 B 2 B
B
#
MAU 3 C H
B U
U S
C H
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DATABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2 Power Supply 1 DC 2
A 2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 NIC 5 (A) (ID = 33) CUSTOM I/O 2
1 B B
PROC 1 = ADA 1, MW 1, UTIL 1, CAL/MCDU 1, CMS 1
U
A 2
AIOPB2
U
#
B
GENERIC I/O 3
1 B
B S
C H
FCM 3
A 1 A 1
PROC 2 = CMF 2 U
#
S
A 1 2 B 2 B
A 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1
2 B 2 B 2 B
Power Supply 2 ESS 2/DC 2 BRAKES (INBD) CONTROL I/O 2 AIOPA2
B S
CUSTOM I/O 1
CONTROL I/O 1 BRAKES (OUTBD) PSEM 1
B
#
Power Supply 3 DC 1 AGM 1
MAU 2
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AFCS 2
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Autopilot System The design of the FGCS relies on a dual channel architecture (i.e., dual redundant AFCS consisting of two AIOP modules per channel) to achieve an automatic fail operational / fail passive capability. Each AFCS channel has a priority manager function, which determines whether a channel is active or stand-by. This determination is based on system / function availability requirements and channel capability. Assuming the channel is capable of performing the intended function, channel priority for all functions, except those performed by the Stall Warning and Protection System, is selectable by the pilot via the MCDU SETUP Page menu. Channel priority is designed in such a way that it automatically transfers system’s operation to the stand-by channel when a failure is detected in the active channel. In order to minimize exposure to latent system failures and to enable the flexibility in the aircraft dispatching (dispatching with only one AFCS channel available), the AFCS is designed in such a way as to automatically transfer the priority after an on-ground power-up to the channel that was designated as the stand-by channel during the previous flight leg, if the other channel (previously being active) is considered failed. The FGCS (AIOP) software modules include processing for FGCS command computations (in accordance with the system’s control laws), as well as processing for the system’s commands and performances monitoring. The parallel servo interfaces with each AFCS channel via a digital bi-directional CAN bus, a motor enable discrete, and a clutch enable discrete. The servo only accepts commands from the master AFCS channel. If more than one AFCS is shown as master indicating a system failure, the servo does not accept an AFCS command. Servo drum position synchro excitation is only provided by the master channel. Servo selection (linkage ratio and PGR ratio) is made on the basis of a predicted control system loads and linkage ratio information provided by the primary flight control system.
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Figure 2: Autopilot System
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AIOP Description
Each AIOP provides its own sensor excitation reference and position sensor input demodulation and filtering for LVDT, RVDT, and resolver type sensors.
The Actuator I/O with Processor Module is a dual slot, single lane module that consists of two interconnected circuit cards (I/O card and Processor card) with an optional mezzanine card. The I/O card contains no software and the processor card communicates directly with the I/O card via the PCI Local bus. The I/O module is primarily dedicated to an actuator control, excitation / demodulation of position sensors, and to the control of functional enable and disable control lines. The I/O card has a generic design and thus has unused I/O capacity for future growth. While the AIOP handles a significant portion of the AFCS I/O, there is AFCS I/O routed through other Primus Epic ® hardware. Signals chosen to be routed through the I/O portion of this card fall into one or more of the following categories:
Analog input processing circuits, including AC demodulation circuits. Demodulated signals are filtered, scaled and they are converted to digital form through a 12 bit A/D.
• Signals whose unique interface requirements (CAN bus, high-current output drivers) are not available on the generic Primus Epic ® I/O cards. • Signals that are used for real-time control and/or monitoring, which require minimization of the system transport lag. • Signals that could be routed through other I/O cards but they are routed through this card because of its spare capacity. The I/O portion of this card contains the following circuits: • • • • • •
ARINC 429 receivers and transmitters Ground – open discrete inputs 28VDC – open discrete inputs Ground – open quick disconnect discrete inputs Low-current 28 VDC - open output discrete High-current 28 VDC - open output discrete (used primarily to interface with servo clutches) The AIOP basic functions can be defined as follows: • AC sensor excitation.
last update: Jun06
• Controller Area Network (CAN) busses. The elevator and aileron smart servos are commanded through this digital serial data bus, and the actuator status is received through the same bus. There are two CAN bus interfaces on the AIOP module. • Heartbeat and power monitor functions that are capable of neutralizing the AIOP discrete outputs. All of the high-current 28VDC-open discrete outputs and selected low-current discrete outputs have additional quick disconnect input hardware logic that preempts processor control and, when activated, will force the output to an open (disengaged) state. The quick disconnect inputs are also read by software as discrete inputs to enable the system disconnect under software control. The architecture of the AFCS/FGCS requires a minimum of two Actuator I/ O with processor modules for engaged operation. The two AIOP’s provide two independent paths for command and monitoring of the autopilot actuators. These modules are identified as lane A and lane B. The two lanes are hardware and software identical, but perform differently, handling separate, complementary, and similar functions depending on the lane. Both AIOP’s (Lane A and Lane B) must be functioning correctly in order for the servos to be actively engaged. The AIOP extracts necessary data from the ASCB bus as well as from directly connected signals. Actuator status, positions and aircraft interface discrete are read and placed for transmission on the ASCB bus. Monitoring of inputs, outputs and internal operations is performed, as necessary, to meet the system safety requirements.
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Figure 3: AIOP Module
last update: Jun06
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System Servos/ Actuators The Fail Passive Smart Servo is a part of the AFCS installed in the Embraer 170. The servo is commanded from the AIOP modules, installed in the MAU cabinets. The two AIOP are used to command each servo via CAN busses, with one AIOP connected to each line of the servo. This configuration allows the AIOP to drive the aircraft flight controls that use the position loop with the software current limit when the actuator is mechanically engaged. The servo provides the information (feedback) to the AIOP about its operational status, using the internal continuous monitoring. The servo architecture is conceived as a dual processor system configuration with functions to drive and monitor the servomotor and to interface/communicate with the AIOPs. The servo actuator consists of the two LRU’s: bracket (mounted in the aircraft) and servo drive. The fail-passive servo consists of the following components: motor, gearbox, DC engage clutch, position sensor and dual independent computation and control circuitry. The servo receives aircraft’s 28 VDC electrical power in addition to its functional interfaces to the AIOPs. The servo provides an interface and the means via hardware to bypass and shutdown the motor drive in each lane. It also provides an interface to allow the AIOP to engage / disengage a clutch between the servo and bracket/ drum assembly mounted in the aircraft. The servo gearbox has interchangeable gears to set torque (and direction) versus speed output characteristics.
last update: Jun06
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Figure 4: SM 3000 SERVO
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Servo Functions and Operations The main servo functions are: • Provide the drum rotation independent from other parts, when the servo drive is not engaged. • Provide a current limited position control when the servo drive is commanded by the AIOP. • Transmit status and feedback information to the AIOP’s when commanded by the AIOP’s. • Provide a resolver feedback of the servo shaft position. The following are servo system operational modes: • • • • •
Power-up Stand-by operational mode Position Loop mode Initiated Test mode (BITE) Data Loop mode
Start-Up Operational Mode.
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Figure 5: SM 3000 Servo Architecture
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Autopilot Aileron Control The Autopilot controls the ailerons using a SM-3OOO Smart Servo. During auto flight the servo responds to commands from the AIOP`s (Actuator input / output Processors) located inside the MAU (Modular Avionics Unit). The Servo is located on the aft torque tube assembly at the centre fuselage. The servo moves the entire Aileron control system through the servo cable wrapped around a capstan. The cable then moves the shafts connected to the Aileron PCU`s. The Hydraulic powered PCU`s then move the ailerons. The control loop is closed by the Aileron Position Sensor sending back the posmon to the AIOP.
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Figure 6: Autopilot Aileron Control
MAU 1
MAU 2
GUIDANCE PANEL
RS-422 CONTROL I/O
CHANNEL A
CHANNEL B
RS-422
A429 AP SELECTION
CONTROL I/O
A429 AP SELECTION AP/FD TCS SW
PILOT MCDU 1
AP/FD TCS SW
COPILOT MCDU 2
TOGA SWITCHES AP QUICK DISCONNECT SWITCHES
SERVO CLUTCH ENGAGE AIOP 1A
AIOP 2A
CAN BUS A CAN BUS B
MAU 3
SERVO POWER ENGAGE
AIOP 1B
AUTOPILOT AILERON SERVO
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AIOP 2B
AUTOPILOT ELEVATOR SERVO
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Figure 7: Aileron
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Aileron PCU’s The PCU’s are mechanically controlled through the cable system and are hydraulically powered. The inboard and outboard PCU’s are powered by hydraulic systems #2 and #3 respectively. The hydraulic systems are distributed as required to provide the redundancy necessary to meet the system safety analysis. The two PCU’s attached to each aileron surface operate in an active configuration. The aileron PCU’s location and attachment are shown. An LVDT is attached to each surface to provide surface position information to the flight data recorder and to the pilots via the EICAS.
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Figure 8: Aileron PCU’s
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Notes:
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Figure 9: Cockpit Transducers
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Hydraulic System (ATA 29-00)
There are three hydraulic systems on the ERJ 170. System pressure for hydraulic systems 1 and 2 is supplied by an engine driven pump (EDP) and an AC electric motor pump (ACMP) for each system. The ACMP’s provide additional hydraulic flow capacity when demand is high and provide backup for the EDP’s following an engine failure. Hydraulic system 3 is powered by two ACMP’s, which are normally on during the entire flight. System 3 also is backed up by the RAT. With the necessity of maintaining roll control with minimal drag for the twoengine-out situation, the outboard aileron PCU on each wing is powered by hydraulic system #3. The inboard PCU on each wing is powered by hydraulic system #2. In the rotor burst area, no single fragment can disable both of the aileron PCU’s on each wing through hydraulic line severance. Hydraulic system #3 is located entirely behind the rotor burst zone and is not affected by such failure. The hydraulic distribution system can be seen.
Autopilot Elevator Control The Autopilot Elevator Servo is located below the control column and connects to the torque tube assembly by cables from the capstan on the servo mount. The Elevator Servo receives signals from the AIOP’s located in the MAU. It will then drive the control column in the desired direction. The four control column transducers (or LVDTs) sense the column movement and send signals to the elevator P-ACE (Powered Actuator Control Electronics). The PACE will then drive the elevator control surface using the elevator hydraulic driven PCU’s (Powered Control Units). The control loop is closed by the PCU LVDT’s and elevator surface LVDT’s sending back elevator position signals to the AIOP’s via the P-ACE and FCM (Flight Control Module).
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Figure 11: Autopilot Elevator Control
MAU 1
MAU 2
GUIDANCE PANEL
RS-422 CONTROL I/O
CHANNEL A
CHANNEL B
RS-422
A429 AP SELECTION
CONTROL I/O
A429 AP SELECTION AP/FD TCS SW
PILOT MCDU 1
AP/FD TCS SW
COPILOT MCDU 2
TOGA SWITCHES AP QUICK DISCONNECT SWITCHES
SERVO CLUTCH ENGAGE AIOP 1A
AIOP 2A
CAN BUS A CAN BUS B
MAU 3
SERVO POWER ENGAGE
AIOP 1B
AUTOPILOT AILERON SERVO
last update: Jun06
AIOP 2B
AUTOPILOT ELEVATOR SERVO
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Elevator Servo An autopilot servo actuator is connected to the left-hand control column on the baseline aircraft. An optional second autopilot servo actuator can be connected to the RH control column. Cockpit Control Transducers (LVDTs) are used to transfer the mechanical motion of the control columns into electrical commands that are sent to the P-ACEs.
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Figure 12: Elevator Servo Installation
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Cockpit Control Transducers The cockpit control transducers convert the mechanical deflection of the control columns into electrical signals that are transmitted to the P-ACES and converted into commands to the hydraulic PCUs to position the elevator surfaces. One LVDT on the pilot’s control column goes to the COM lanes on the two P-ACEs for the left elevator and the other LVDT goes to the MON lanes on the same P-ACEs. Similarly, one LVDT on the co-pilot’s control column goes to the COM lanes of the two P-ACEs for the right elevator and the other LVDT goes to the MON lanes on the same P-ACEs.
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Figure 13: Elevator Architecture
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Hydraulic System Interface There are three hydraulic systems on the ERJ 170. System pressure for hydraulic systems 1 and 2 is supplied by an engine driven pump (EDP) and an AC electric motor pump (ACMP) for each system. The ACMPs provide additional hydraulic flow capacity when demand is high and provide backup for the EDPs following an engine failure. Hydraulic system 3 is powered by two ACMPs, which are normally on during the entire flight. System 3 also is backed up by the RAT. The inboard elevator PCUs are on hydraulic system #2. The left outboard PCU is on hydraulic system #1. The right outboard PCU is on hydraulic system #3. In the rotor burst area, no single fragment can disable all of the elevator system PCUs through a combination of hydraulic lines or electrical harness severance. Hydraulic systems #1 and #3 are located entirely behind the rotor burst zone and are not affected by such failure.
Rudder Turn Coordination The AFCS supplies the rudder turn coordination to eliminate the excessive centripetal or centrifugal forces during a banked turn. The AIOP sends rudder turn coordination commands through the ASCB bus to the FCMs. The FCM sends rudder turn coordination data to the ACE on the CAN bus. The AIOP receives from the FCM a verification of its command through the ASCB. The loss of the rudder turn coordination function does not cause the yaw damper to disengage.
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Figure 15: Rudder Turn Coordination Signal Flow
ASCB AIOP1A
FCM 2 CAN BUS
AIOP1B
P-ACE
FCM 4
Upper Rudder PCU
CAN BUS
SURFACE SENSOR
CAN BUS
AIOP2A
FCM 1
P-ACE
Lower Rudder PCU
R U D D E R
SURFACE SENSOR
CAN BUS
AIOP2B
FCM 3 Filename: Rudder Control Signal Flow 6/13/2002 Rev -
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Autopilot Rudder Control (for CAT III) The Autopilot Rudder Servo is located below the control column and connects to the torque tube assembly of the rudder pedals, by cables from the capstan on the servo mount. The Rudder servo receives signals from the AIOP’s located in the MAU. It will then drive the Rudder pedals in the desired direction. The rotation of the torque tube is sensed by four Rudder pedal transducers (or LVDTs). These send signals to the Rudder P-ACE (Powered Actuator Control Electronics). The P-ACE will then drive the Rudder control surface using the Rudder hydraulic driven PCU’s (Powered Control Units). The control loop is closed by the PCU LVDT’s and rudder surface LVDT’s sending back Rudder position signals to the AIOP’s via the P-ACE and FCM (Flight Control Module).
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Figure 16: Autopilot Rudder Control
MAU 1
MAU 2
G UIDANC E P ANEL RS4 - 22 C ONTR OL I/O
C HANNEL A
C HANNEL B
RS4 - 22
A429 AP S ELEC TION
C ONTR OL I/O
A429 AP S ELEC TION AP/FD T CS S W
P ILOT MCDU 1
AP/FD T CS S W
C OP ILOT MCDU 2
T OG A S WITC HE S AP QUICK DIS CONNEC T S WITC HE S
SERV O C LUTC H E NG AG E AIOP 1A
AIOP 2A
C AN B US A C AN B US B
MAU 3 SERV O P OWER E NG AG E
AIOP 1B
AUT OP ILOT AILER ON SERV O
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AIOP 2B
AUT OP ILOT R UDDER SERV O
AUT OP ILOT E LE VATOR SERV O
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Optional Rudder Servo (CAT III a) Optional autopilot servo actuators may be attached to each torque tube assembly to provide autopilot input into torque tube assemblies and the pedal assemblies. As with normal pedal operation and trim input, autopilot input will result in rudder commands from the command LVDTs attached to the forward torque tube assemblies. The autopilot servo control of the pedal input is shown.
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Figure 17: Rudder Servo (optional CATIIIa)
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Cockpit Control Transducers The cockpit control transducers convert the mechanical deflection of the pedals into electrical signals that are transmitted to the P-ACEs and converted into commands to the hydraulic PCUs to position the rudder surfaces. Four single LVDTs are provided as cockpit control transducers. Two are attached to the captain’s pedal tower and two to the first officer’s pedal tower. Each LVDT has one end grounded to structure and the other end attached to a separate bell crank on the appropriate tower. The bell crank is equipped with a shear-out to protect the rudder pedals from jamming in the unlikely event of a jammed LVDT. In this arrangement, two of the transducers are commanded by the pilot and two by the co-pilot.
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Figure 18: Rudder control schematic
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Hydraulic System Interface There are three hydraulic systems on the ERJ 170.System pressure for hydraulic systems 1 and 2 is supplied by an engine driven pump (EDP) and an AC electric motor pump (ACMP) for each system.The ACMPs provide additional hydraulic flow capacity when demand is high and provide backup for the EDPs following an engine failure. Hydraulic system 3 is powered by two ACMPs, which are normally on during the entire flight.System 3 also is backed up by the RAT. The inboard elevator PCUs are on hydraulic system #2.The left outboard PCU is on hydraulic system #1.The right outboard PCU is on hydraulic system #3.In the rotor burst area, no single fragment can disable all of the elevator system PCUs through a combination of hydraulic lines or electrical harness severance.Hydraulic systems #1 and #3 are located entirely behind the rotor burst zone and are not affected by such failure.
• The aileron or elevator control system disconnect unit shows that the pilot and copilot control systems are no longer connected.
AUTOPILOT SYSTEM The autopilot system supplies the pitch and roll control with the autopilot servos and aircraft trim system. The AFCS pitch and roll commands to the autopilot servos go through the CAN bus. The autopilot control authority has the constraints that follow:
• A column force monitor trips.
• Limits are set so that the maximum overpower forces are not more than 75 lb in pitch and 50 lb in roll, as measured at the control column. • The aircraft roll rate has a limit of +/- 7.5 degrees or second. • The aircraft pitch rate limit is a function of the true airspeed to limit the acceleration changes to +/- 0.3 g in a straight and level flight. • The autopilot can engage through the range of +/- 25 degrees pitch and +/- 35 degrees roll. When engaged, the autopilot system decreases the pitch and roll angles below those control limits. • The autopilot control limit for roll is +/- 35 degrees unless it is in the APR (approach) mode. • The limits decrease from +/- 25 degrees above 200 feet radio altitude to +/- 5 degrees at 0 ft radio altitude when in the APR mode. • The autopilot pitch control limit is 20+/- degrees. In the APR, the nosedown pitch limit is programmed below 300 ft. The autopilot is set with the autopilot engage/disengage push button on the GP. When the autopilot is engaged, the autotrim function and YD (Yaw Damper) function start if they are not already on. The manual cancellation of the autopilot system does not cancel the YD. The autopilot system disengages when: • • • •
The autopilot system can engage with or without active FD guidance modes. When no vertical FD mode is on, the autopilot system is in the flight path angle (FPA) hold mode. When no lateral AP mode is on, the autopilot system is in roll hold, wings level, and/or heading hold. A YD failure with the autopilot engaged does not cause the autopilot system to disengage. When pitch or roll FD guidance modes have been set, the autopilot system connects itself to the pitch and/or roll commands given by the flight guidance function. When the TCS button on the pilot or copilot control yoke is pushed, the AFCS releases the aileron and elevator servo clutches to neutralize the autopilot system. The release of the TCS button causes the autopilot clutches to re-engage. The letters AP come into view on the PFD when the autopilot system is engaged. The AP annunciator on the PFD shows when the autopilot is engaged. The autopilot annunciation is on the FMA section of the PFD. When the autopilot is disengaged normally, an aural alert is started. When the autopilot disengages abnormally, a caution CAS (Crew Alerting System) message and an aural alert are given. To cancel the annunciation and the aural alert, the pilot or copilot pushes and holds the AP/TRIM DISC push button on the control yoke for a period of 2 s minimum.
The manual trim switches are activated. The QD switches are activated. The stick shakers are activated. The fly-by-wire system is in the direct mode.
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Figure 20: Flight Guidance and Control System
CONTROL YOKE
CONTROL YOKE
AILERON
ELEVATOR
RUDDER PEDALS ELEVATOR SERVO
LVDT
MAU'S
LVDT
AIOPS
FCMS
P-ACES
AILERON SERVO
ACTUATOR
ACTUATOR
ACTUATOR
ACTUATOR
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HYDRAULIC
CAN BUS
MECHANICAL
ELECTRICAL
ACTUATOR
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CAT (category) II (APPR II) Engagement The CAT II logic is automatically activated when the RA/BARO minimums selector knob is set to the RA position. A green APPR 2 annunciator on the FMA indicates the correct setting and a white or amber APPR 1 ONLY annunciator indicates an incorrect setting. CAT II operation is allowed only with the green APPR 2 annunciator enabled and the AP coupled. To get a green APPR 2 annunciator the following conditions must be met: • Radio altitude below 1500 ft. • Flaps 5. • NAV (Navigation) 1 on pilot side and NAV 2 on copilot side, both NAVs tuned to the same LOC (Localizer) frequency. • An active approach GS/LOC mode selected. Both courses set to the same value. • Both flight directors operational. • Attitude and heading valid on both PFDs. • Glide slope and localizer deviation valid on both PFDs. • No reversions (IRS and ADC (Air Data Computer)) modes selected on both PFDs. • Valid airspeed and baroaltitude on both PFDs. • No comparison monitors are tripped (FPA, attitude, heading, airspeed, baroaltitude, localizer, glide slope and radio altitude) on both PFDs. • No back course selected. • The CAS message APPR 2 NOT AVAIL not present. • CAT II decision height setting on both display control panels (greater than 80 ft and less than 200 ft valid and set to RA). • RA/BARO minimums selector knob set to RA. • No TCS button pressed.
If the green APPR 2 annunciation is displayed and one of the following conditions is achieved, the amber APPR 1 ONLY annunciation flashes active characters inverse video for 5 seconds, then steady in conjunction with RA minimum selected digital read-out: • • • • •
No valid radio altitude displayed. Aircraft no longer APPR 2 capable. Flaps position other than below 800 ft. CAS message SLAT-FLAP LEVER DISAG shows. Either minimums selected readouts change from RA to BARO.
LOC frequency or inbound course mismatch.
When the green APPR 2 annunciator is enabled, the localizer lateral deviation scale is expanded with the external limits representing the excessive deviation points. last update: Jun06
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Figure 21: Approach Sequence CAT I, CAT II
Approach Sequence – CAT II (APPR 2 Available)
Approach Sequence CAT -I (RA/BARO SET to RA)
Approach Sequence CAT I - (APPR 2 Not Available)
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
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170/190 MAINTENANCE TRAINING MANUAL
ILS APPROACH During execution of the ILS approach, Autopilot Approach Status Annunciator displays the current status of the system and alerts whether the intended approach matches system capabilities. The RA/BARO selector and RA Minimums setting inform the system what is the intended approach. When ILS modes are requested via APP button, system arms for the highest capability available. If all necessary requirements are not accomplished, an EICAS message is presented during flight and informs that category II ILS approach mode is not available. The intended approach is informed to the system setting the barometric correction via control knobs on Display Controller panel (guidance panel). • CAT1 – set RA/BARO selector to BARO (both sides) • CAT2 – set RA/BARO selector to RA and adjust Minimums to 80 ft or above The operational conditions to accomplish a CAT II approach are: • • • • •
RA/BARO set to RA and Minimums set at 80 ft or above Both NAV set to correct LOC frequency Both PFDs set to correct LOC inbound course (V/L or Preview) Flap 5 All described conditions established at or above 800 ft RA
If the flap setting is the only remaining condition to be satisfied for CAT II, the armed status will remain displayed down to 800 ft RA, suggesting there is still one pilot’s action pending. The ILS approach check points are the following: • 1500 ft RA – system starts trying to engage highest capability available. 800 ft RA – system “freezes” highest capability available, not allowing approach “upgrades” anymore. last update: Jun06
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Chapter 22-10
Page 53
Figure 22: Approach Sequence CAT I, CAT II
Approach Sequence – CAT II (APPR 2 Available)
Approach Sequence CAT -I (RA/BARO SET to RA)
Approach Sequence CAT I - (APPR 2 Not Available)
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
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170/190 MAINTENANCE TRAINING MANUAL
Autoland CAT IIIa, CAT IIIb (Rudder Servo Installed) Definitions • Fail Operational: A Fail Operational System is a system which after failure of any single component, is capable of completing an approach, flare and touchdown, or approach, flare, touchdown and rollout by using the remaining operation elements of the Fail Operational System. • Fail Passive: A Fail Passive System is a system which in the event of a failure, causes no significant deviation of aircraft flight path or attitude. the capability to continue the opperation may bo lost and an alternate course of action (e.g., a missed approach) may be required.
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 55
Figure 23: Definitions Fail-passiv Autoland refers to the following:
Fail-Passive Autoland
System Capability
PVR
DH
Operational Approval
AUTOLAND 1 Fail-Passive Autoland No Rollout Guidance
600 ft (175 m) RVR TD 400 ft (125 m) RVR TD 400 ft (125 m) RVR RO
Not less than 50 ft HAT
CAT IIIb
last update: Jun06
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Chapter 22-10
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170/190 MAINTENANCE TRAINING MANUAL
Autoland
Rollout
Autoland 1 • Fail-Passive Autoland with Limited Rollout capability
The ailerons are gently moved toward neutral during ground rollout mode. This allows the aircraft to maintain a safe lateral trajectory until the nose lowering control law has completed the landing.
Autoland 2 • Fail-Passive Autoland with Extended Rollout capability
Vertical guidance to bring the nose gear into contact with the runway.
Unique Autoland Modes: – – – –
De-Rotation
Align Flare Rollout De-Rotation
Align Mode Magnetic Heading and Track Angle are used to produce guidance through a parallel rudder servo to align the aircraft with the runway prior to touchdown, reducing any “crab” angle due to cross wind.
Flare Vertical guidance to transition from glideslope track to touchdown, landing the aircraft within the longitudinal runway dispersion requirements. A flare bias is injected into the horizontal stabilizer trim command during Autoland and prior to engagement of the flare mode. It produces a definite, but slight, nose up trim in order to prevent a nose down transient in the event of an automatic disconnect during flare.
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 57
Figure 24: Autoland Sequence
AUTOLAND 2 1500ft AUTOLAND 2 SPD RETD AP T ALIGN 150ft AT AT FLARE 50ft RETD 30ft Main Wheel Touchdown AT quick disconnect button Touchdown +5seonds (Autoland AP 1)
Autoland Failures Failure response when Autoland is Engaged varies with height. Above 800 ft automatic channel transfers may allow AP and Autoland to remain engaged. Below 800 ft automatic channel transfers result in loss of autopilot or Autoland depending on the failure. Autoland 2 is not allowed to degrade directly to Autoland 1
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-10
Page 59
Figure 25: Autoland failures - FLAPS NOT SET TO 5
1500 ft Autoland ist armed but unable to engage because the flaps are not set to 5. Aircraft secends below 800 ft RA Minimums Box also flashes amber as thes needs to be re-set to CAT 1 BARO MIN
AUTOLAND 2 APPR 11 ONLY APPR ONLY SPDT AP
AT APPR 1 ONLY APPR 1
RA 50
APPR 1 ONLY disapears when minimums correctly set APPR 1
OR GO AROUND
last update: Jun06
APPR APPR 11 LOC GS
FOR TRAINING ONLY - Reproduction Prohibited
BARO RA 200 50
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Autoland Failures If the autopilot remains engaged, loss of Autoland results in a red "NO AUTOLAND" annunciation being displayed and if appropriate an amber flashing minimums box on the PFD. The modes ALIGN, FLARE, D-ROT, and RLOUT are removed from the FMA and a single "NO AUTOLAND" aural alert sounds. The red "NO AUTOLAND" annunciation is removed when any of the following is true: • The TOGA pushbutton is pressed • The AP quick disconnect pushbutton on either wheel is pressed • The FGCS flight director mode is Go-Around or Takeoff Autopilot warnings have priority over Autoland warnings
last update: Jun06
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Chapter 22-10
Page 61
Figure 26: Autoland failures - REVERSION TO APPR2
RALT 1 FAIL S ystem re verts to highest a va ilable approach mode.
NO AUTOLAND APPR 2 ONLY SPDT AP AT
APPR 2 LOC GS ALIGN
FLARE
After 5 seconds the AP PR 2 ONLY caution appears to prompt crew. When the correct CAT 2 minimums is s et the caution disappea rs. OR GO AROUND last update: Jun06
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Notes:
last update: Jun06
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Chapter 22-10
Page 63
Figure 27: Autoland failures - AUTOPILOT DISC
AP FAIL (inhibited below 200’) AP QUICK DIS C
SPDT
AP AT
AUTOLAND 2 ROLL ROLL FLARE FPA ALIGN FPA RLOUT
D-ROT
MANUAL LANDING
RA 50
last update: Jun06
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22-11 Guidance and Display Guidance Panel A single Guidance Panel (GP) is dedicated to provide the means for selection and/or de-selection of most of the AFCS functions. Additional functions related to display control are also included within the GP, as a part of the left and right DCP. The GP is conceived in such a way that it contains the two identical independent channels, each providing communication to one AFCS channel (A or B). In addition, isolation of electronics, switch contacts (isolated with diodes), and dual power supplies are utilized within the GP in order to prevent a single fault from affecting both AFCS channels. After pressing any GP functional push button to arm a mode, a second press of the same push button will disarm that mode. Annunciations of the operational modes and autopilot engagement status are also provided on the PFD in the FMA area. The other FGCS annunciations and indications, related to the FGCP function / operation are presented on the appropriate MCDU pages. The flight guidance control system indirectly drives the primary flight controls, through either the cockpit control column or through other processing modules.
last update: Nov06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 22-11
Page 1
Figure 1: The automatic flight control system
Diplay Controller
last update: Nov06
Guidance Control
FOR TRAINING ONLY - Reproduction Prohibited
Diplay Controller
Chapter 22-11
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170/190 MAINTENANCE TRAINING MANUAL
Guidance Panel The GP is on the glareshield panel in the cockpit. The dual channel GP supplies the means for the selection of all FGCS flight director functionality. Each channel communicates with one AFCS channel 1 or 2. For the FGCS functions, the GP serially transmits the pushbutton data through the private RS-422 serial links to the MAU control I/O modules. The FGCS uses a second path (direct discretes) to send the GP a validation of the operator selections by turning the annunciators for each pushbutton on or off. If an FD mode is armed or engaged, the annunciator for the pushbutton comes on. The annunciation of the modes and the autopilot engage status also shows on the PFD. The controls on the GP are placed in five functional groups as follows: • • • • •
Lateral guidance control AFCS control Speed control Vertical guidance control DCP (Display Control Panel)
last update: Nov06
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Chapter 22-11
Page 3
Figure 2: Flight Guidance and Control System (FGCS)
LEFT DC
LEFT GP
(J1)
(J2)
ON-SIDE IRS REV PB
ON-SIDE ADC REV PB
CROSS-SIDE IRS REV PB
AP QD RIGHT GP
CROSS-SIDE ADC REV PB
COPILOT CONTROL YOKE TCS NO. 2
TCS NO. 1
AP QD
PILOT CONTROL YOKE
TOGA SIGNAL 2
THROTTLE CONTROL QUADRANT (TCQ) TOGA SIGNAL 1
CROSS-SIDE ADC REV PB
T/O CONFIG CHECK PB
CROSS-SIDE IRS REV PB
EICAS DECLUTTER OVERRIDE PB
ON-SIDE IRS REV PB
PILOT CHRONOMETER PB
ON-SIDE ADC REV PB
REVERSIONARY PANEL
COPILOT CHRONOMETER PB
RIGHT DC
GUIDANCE PANEL (J3)
(J4) MAU 2
MAU 1
(RS-422) (RS-422)
BACKPLANE BUS
last update: Nov06
(RS-422)
ANNUN VALID
ANNUN VALID
BTN PNL ARM
BTN PNL ARM
DATA (RS-422)
DATA (RS-422)
DATA (RS-422)
DATA (RS-422)
STROBE (RS-422)
STROBE (RS-422)
CLK (RS-422) NIC
ASCB
CONTROL I/O 2
CLK (RS-422) TO OTHER AIRCRAFT SYSTEMS
TO OTHER AIRCRAFT SYSTEMS
FOR TRAINING ONLY - Reproduction Prohibited
ASCB
NIC
BACKPLANE BUS
(RS-422) CONTROL I/O 1
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170/190 MAINTENANCE TRAINING MANUAL
22-17 AFCS Warnings and Indications Figure 1: FMA (Flight Mode and Annunciation)
Basic FMA will include five boxes depicting:
SPEED/AT MODE
STATUS
SRC
LATERAL
VERTICAL
Lateral/Vertical FD modes will be combined when applicable ( ie. Approach modes) Armed FD and SPEED modes will be displayed below active line When the FMS source is selected the FD mode will be identified by color without the use of the letters V and L
AFCS Channel Select on MCDU The MCDU gives the option to select the channel that the AFCS uses.The SETUP page shows this option.The LSK (Line Select Key) 3R changes between channel A and channel B.The selected channel shows in larger size letters.To access the SETUP page, the pilot or copilot pushes the MENU key on the MCDU.Then the pilot or copilot pushes the LSK 1L for the MISC MENU.On the MISC MENU, the pilot or copilot pushes the LSK 2L to show the SETUP page.
Figure 4: Autopilot Channel Selection
The Auto P ilot ha s two cha nne ls which a lte rna te a utoma tica lly by s ys te m or by pilot input.
The AP provides automatic pitch and roll control of the airplane through the elevator and aileron AP servos.
last update: Nov06
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170/190 MAINTENANCE TRAINING MANUAL
AP/TRIM Disc Pushbutton The AP/TRIM disconnect pushbutton on the pilot and copilot control yokes is a momentary switch that does a QD (Quick Disconnect) of the autopilot functions.The switch supplies an output to the GP and to the AIOP modules. The GP transmits the state of the switch to the ASCB through the control I/ O 1 and 2 modules.The AIOP modules sense the state of the switch and make the servo clutch drives unserviceable. When the autopilot disengages, a reverse video flashing AP (Automatic Pilot) shows on the FMA (Flight Mode Annunciator) area on the PFD.Likewise, when the autopilot disengages, the pilot and copilot hear an “AUTOPILOT” aural warning.The AP/TRIM disconnect pushbutton lets the pilot or copilot turn off the AP annunciator and aural warning.The annunciator and aural warning turn off when the pilot or copilot pushes the AP/TRIM disconnect pushbutton and holds it for more than two seconds.
AP/FD TCS Pushbutton The AP/FD TCS pushbutton makes the autopilot servo clutches disengage. When the TCS (Touch Control Steering) pushbutton is released, the autopilot system automatically engages the servo clutches. When the TCS pushbutton is pushed during takeoff, and the aircraft is below 400 feet AGL (Above Ground Level), the FGCS disengages the takeoff mode when the aircraft reaches 400 feet AGL.
Figure 5: Autopilot Dis-engagement
Manual Trim switch
TCS
Quick Disconnect The autopilot shall command the servo clutches to disengage while TCS is active. The autopilot shall automatically reengage the servo clutches and resynchronize references when TCS deactivates. last update: Nov06
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Notes
Figure 6: CAS Messages
CAS MESSAGE
DESCRIPTION
FAULT CODE
ADVISORY
Loss of Redundancy of the AFCS
22102200
AFCS PANEL FAIL
ADVISORY
Failure of the AFCS portion of the Guidance Panel
22100200
AFCS PANEL FAULT AP FAIL AP FAULT AP PITCH MISTRIM AP PITCH TRIM FAIL
ADVISORY CAUTION ADVISORY CAUTION CAUTION
22100300 22100400 22100500 22100600 22100700
AP PITCH TRIM FAULT
ADVISORY
AP ROLL MISTRIM APPR2 NOT AVAIL
CAUTION ADVISORY
Fault of the AFCS portion of the Guidance Panel Failure of the Autopilot Failure of a Single Channel of the Autopilot The AP Detects a Pitch Mistrim Condition Failure of the AP Pitch Trim Function Failure of a Single Channel of the AP Pitch Trim Function The AP Detects a Roll Mistrim Condition CAT II Is Not Available Failure of the Autothrottle Function Failure of a Single Channel of the Autothrottle Function The Autothrottle is not in Hold
AFCS FAULT
AT FAIL AT FAULT
last update: Nov06
TYPE
CAUTION ADVISORY
22100800 22100900 22000100 22300100 22300200
AT NOT IN HOLD
CAUTION
DISPLAY CTRL FAIL
CAUTION
DISPLAY CTRL FAULT
CAUTION
ENG TLA NOT TOGA
CAUTION
ENG TLA TRIM FAIL
ADVISORY
Failure of Engine TLA Trim
22300500
FD FAIL
ADVISORY
Failure of Flight Director
22101100
FD FAULT
ADVISORY
Failure of a Single Channel of the Flight Director
22101200
FD LATERAL MODE OFF FD VERT MODE OFF
Failure of the Flight Director Lateral Mode Failure of the Flight Director Vertical Mode
22101300 22101400
YD FAIL
CAUTION CAUTION ADVISORY
Failure of the Yaw Damper
22101700
YD FAULT
ADVISORY
22101800
YD OFF
ADVISORY
Failure of a Single Channel of the Yaw Damper Yaw Damper is Off
Failure of the Display Controller portion of the Guidance Panel Fault of the Display Controller portion of the Guidance Panel The Thrust Lever is not in TOGA Position
FOR TRAINING ONLY - Reproduction Prohibited
22300300 22102000 22102100 22300400
22101900
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Auto Pilot System Diagnostic Tests SYSTEM DIAGNOSTICS MENU Select System Diagnostics in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the System Diagnostics Soft Key • Select the System Diagnostics Soft Key by pushing one of the enter keys on CCd No.2 • The System Diagnostics menu is displayed and a list of member systems organized by ATA chapter that have system diagnostic pages associated with them are presented.
Figure 7: Autopilot System Diagnostics Tests
MAINTENANCE M ESSA GES
MEM SYSTEM BER DA SYSTEM IA GNOSTICS SATUS
EXTENDED MAINTENA NCE
DATA LOA DER
FILE TRANSFER
last update: Nov06
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Figure 8: Autopilot System Diagnostics Tests
Intentionally left blank
last update: Nov06
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22-30 Auto thrust Thrust Management System Definition System Definition Overview The ERJ 170 TMS is configured in a dual redundant architecture with dual redundant TRS, ETTS and AT. The dual configuration of the architecture is designed for increased system availability. Only one channel of the TMS, only a single TRS, ETTS and AT is operating at any given time. The pilot can select the priority AT channel as well as TRS and ETTS channels via the MCDU.
Figure 1: System Overview MAU
MAU
LE FT ENGINE FADEC
A RINC 429 FADEC to MAU ARINC 429 MAU to FADEC CH A
ARINC 429 MAU to FADEC CH A
Channel A
Channel B
Generic I/O Module 1
ARINC 429 FADEC to MAU Generic I/O Module 3
RIGHT ENG INE FADEC ARINC 429 FADEC to MAU
Channel A ARINC 429 MAU to FADEC CH B
A RINC 429 MAU to FADEC CH B Channel B
AIO / Proc Module 1A y y y y
ETTS Thrust Rating Selection AT Control Law AT Engage/ Disengage Logic
AIO / Proc Module 1B y y
AT Disengage Logic AT T/O Clamp Monitor
ARINC 429 MAU to TCQ
ARINC 429 FADEC to MA U
Thrust Control Quadrant (TCQ ) Left A T Servo
ARINC 429 TCQ to MAU
Right AT Servo
ARINC 429 MAU to TCQ
ARINC 429 TCQ to MAU AT Quick Disconne ct Signal 2
AT Quick Disconnect Signal 1 Left AT Servo Enable
Left AT Servo E nable
Right AT Servo Enable
AIO / Proc Module 2A y y y y
ETTS Thrust Rating Selection AT Control Law AT Engage/ Disengage Logic
AIO / Proc Module 2B y y
AT Disengage Logic AT T/O Clamp Monitor
Right AT Servo Enable
Guidance Panel Control I/O Module 1
GP Serial Bus AT Engage/Disengage Switch
GP Serial Bus A T Engage/Disengage Switch
Control I/O Module 2
ARINC 429 - MCDU Output B us MCDU 2
ARINC 429 - MCDU Input Bus
ARINC 429 - MCDU O utput Bus A RINC 429 - MCDU Input Bus
MCDU 1
ASCB-D NIC
NIC To Other Aircraft Systems
last update: Nov06
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170/190 MAINTENANCE TRAINING MANUAL
System Function Description The Thrust Management System for the ERJ-170/190 is comprised of the following 3 functions:
Auto throttle system (AT) The Auto throttle System uses data from the AFCS Flight Director and Flight Management systems to provide automatic, full flight regime energy management with a minimum of pilot inputs. Flight economy is improved by accurate speed control and thrust management. Safety is enhanced by maintaining aircraft speed and engine speed/thrust within the minimum/ maximum operating limits which helps to reduce pilot workload. AT system operation is the same as previously certified Primus 2000 AT systems.
Electronic Thrust Trim System (ETTS) The ETTS assists the pilot or auto throttle system to match thrust response from each engine, as well as to synchronize N1 or N2 engine fan speeds. System engage/disengage status is pilot selectable. ETTS system operation is the same as previously certified Primus 2000 AT systems.
Thrust Rating Selection function The Thrust Rating Selection function determines the thrust rating (e.g. MCLB, MCRZ, MCT, MTO) based on the phase of flight and pilot entry (for manual rating selection). The active upper engine rating is displayed on the EICAS display N1 dials. TRS system operation is the same as previously certified Primus 2000 FMS systems.
Figure 2: AT Function Description
AT Disconnect
AUTO THROTTLE ENGAGE BUTTON ON GUIDANCE PANEL
TOGA switch
last update: Nov06
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170/190 MAINTENANCE TRAINING MANUAL
Auto throttle System (AT) The baseline ERJ170 contains a dual Auto throttle System. The Auto throttle System uses data from the FADEC, TCQ, MCDU, AFCS Flight Director, and Flight Management systems to provide automatic, full flight regime energy management with a minimum of pilot inputs. Flight economy is improved by accurate speed control and thrust management. Safety is enhanced by maintaining aircraft speed and engine speed/thrust within the minimum/ maximum operating limits which helps to reduce pilot workload. The Auto throttle will automatically position the throttle levers for speed or thrust control by closing the loop on Throttle Resolver Angle from the FADECs. Only the high priority Auto throttle will perform this function. The following features are provided by the Embraer ERJ170 Auto throttle system: • The dual TMS allows either AFCS channel to control the throttle servos and ETTS. Priority automatically switches when the high priority AT channel detects a failure and the other channel is still capable. If the AT was engaged when this priority transfer occurs, it is automatically disengaged and must be manually re-engaged by the crew. Manual channel selection of the AT is available to the flight crew via the MCDU, provided the two channels are equally capable. • Full Flight Regime Auto throttle Operation including: • Takeoff • Climb • Cruise • Approach (Glides lope) • Go Around • Retard • Winds hear • FADEC will provide digital parameters such as Engine Ratings (TO, GA, CON, CLB1, CLB2, CRZ),
Figure 3: PFD- Mode Annunciators
Autothrottle Annunciation
- AT disengaged normally
- AT disengaged abnormally
- AT override function is active & AT is replaced by OVRD last update: Nov06
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170/190 MAINTENANCE TRAINING MANUAL
Electronic Thrust Trim System (ETTS) The ETTS assists the pilot or Auto throttle System to match thrust response from each engine by synchronization of N1 and engine trim. System engage/ disengage status follows Auto throttle engagement status. The Electronic Thrust Trim System (ETTS) commands limited authority thrust changes for the FADECs and outputs the commands on ASCB. The Generic I/O module subsequently reads the commands from the ASCB and transmits them to the FADECs via ARINC 429 busses. The trim system performs TRA trimming to assist the pilot and AT by equalizing the thrust setting for both engines. When the throttles are within trim authority of the active N1 rating, and the auto throttle is not engaged, the ETTS will trim the engines to achieve the active N1 rating. The FADEC transmits to the TMS the trim authority available to the ETTS.
Figure 4: Electronic Trim System
last update: Nov06
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Thrust Rating Selection Function The Thrust Rating System provides automatic selection of the appropriate thrust (N1) rating based on the current phase of flight. The selected rating is displayed on the EICAS. The MCDU provides a means for manually entering a thrust rating, and selecting the AT source and ETTS mode. The MCDU also provides a means for the crew to select the priority channel for the thrust rating function, provided the two channels are equally capable. The Thrust Rating Selection (TRS) function determines the thrust rating (e.g. TO, GA, CON, CLB1, CLB2, CRZ) based on the phase of flight and pilot entry (for manual rating selection). The active upper engine rating is displayed on the EICAS display N1 dials. The Embraer ERJ170 Thrust Management System contains the following components: • Thrust Management System software which resides and executes on the AIO modules, • An ARINC 429 interface with the Thrust Control Quadrant (TCQ), • ARINC 429 interfaces (located on the generic I/O modules) with both FADECs, • Two AT servos located in the TCQ, • One AT engage/disengage switch located on the Guidance Panel, • Two AT quick disconnect switches located on the throttle levers, • An AT source selection on the MCDU, and • An engine rating selection for TRS on the MCDU.
Figure 5: Thrust Rating Selection Function
Provides automatic selection of the appropriate N1 rating based on the phase of flight THRUS T RATING Provides input for display S ELECTION ON Allows for manually selected thrust ratings MCDU Thrust ratings include: MTO MCLB MCRZ Manual
last update: Nov06
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Figure 6: TMC CAS Messages
#
Cat.
Message
1
A
AT FAIL
2
A
AT FAULT
Description
3
C
AT NOT IN HOLD
4
A
ENG SYNC FAIL
5
A
ENG SYNC LIMIT
last update: Nov06
CASCADED MESSAGE
INHIBITED FLIGHT PHASE
Both AT Channels are failed.
K3
A Single AT Channel failed.
K3 K4 K5
Autothrottle is not in TO Hold while the aircraft is in TO Hold mode r efion (beyond 60 kt IAS during TO roll and below 400 ft AGL) Selected Sync function is unavailable Selected Sync function is unable to perform the pilot selected function to an authority limit
FOR TRAINING ONLY - Reproduction Prohibited
K3 K5 K3 K5
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22-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 22-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 22 AUTO FLIGHT ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 10-00 Autopilot ¦ Channels ¦
B ¦ 2 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ operations do not require their ¦ use.
¦ ¦ ¦
¦ 10-12 Mach Trim ¦ Channels ¦ (ERJ-190)
B ¦ 2 ¦ ¦
¦ 1 ¦ ¦
¦ ¦ ¦
| ¦ | ¦ | ¦
¦ 10-14 Yaw Damper ¦ Channels
B ¦ 2 ¦
¦ 1 ¦
¦ ¦
¦ ¦
¦ 10-16 Flight Director ¦ Channels ¦
B ¦ 2 ¦ ¦
¦ 1 ¦ ¦
¦ One channel may be inoperative ¦ provided operations do not require ¦ its use.
¦ 10-20 Takeoff/Go-Around C ¦ 2 ¦ (TO/GA) Buttons ¦
¦ 1 ¦
¦ ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 22-2 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 22 AUTO FLIGHT ¦ ¦ ¦
¦ 11-01 Guidance Panel ¦ (GP)
¦ ¦
¦ ¦
¦ ¦
¦
1) GP Channels
C ¦ 2
¦ 1
¦
¦
¦ ¦ ¦
2) Flight Director (FD) Buttons
C ¦ 2 ¦ ¦
¦ 0 ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦
3) Autopilot (AP) B ¦ 1 Button ¦
¦ 0 ¦
¦ May be inoperative provided ¦ operations do not require its use.
¦ ¦
¦ ¦ ¦
¦ ¦
4) Yaw Damper (YD) Button
C ¦ 1 ¦
¦ 0 ¦
¦ ¦
¦ ¦
¦ ¦
¦ ¦
5) Source (SRC) Button
C ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided ¦ operations do not require its use.
¦ ¦
¦ ¦ ¦ ¦
6) Airspeed to Mach (PUSH IAS/MACH) Change Button
C ¦ 1 ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦
7) Navigation (NAV) Mode Button
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ procedures do not require its use. ¦
¦ ¦ ¦
¦ ¦
8) Heading (HDG) Mode Button
B ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ ¦
¦ ¦
9) Approach (APP) C ¦ 1 Mode Button ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ One may be inoperative on non¦ flying pilots side provided: ¦ a) Autopilot is not used at ¦ less than initial approach ¦ altitude, and ¦ b) Repairs are made within two ¦ flight days.
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
16) Feet to Meter (PUSH FT-M) Change Button
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ operations do not require its use. ¦
¦ ¦ ¦
¦ 11-03 Autopilot/Flight ¦ Director Touch ¦ Control Steering ¦ (AP/FD TCS) ¦ Buttons
C ¦ 2 ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
17) Flight Path Angle (FPA) Mode Button
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ procedures do not require its use. ¦
The VHF Communications system provides two-way air-to-air and air-toground communication in the frequency range of 118.000 to 136.975 MHz with 8.33 or 25 kHz channel spacing which is selectable by the flight crew. The system has an automatic transmit time-out algorithm to prevent blockage of a communication channel if a mic/boom PTT is stuck active for some reason. An optional third VHF Communication system, as a separate LRU and antenna can be installed in the aircraft. The third VHF Comm provides voice and data capability.
• and the Passenger announcement system.
The VHF COM system interfaces to following aircraft subsystems:
The ERJ170 communication system is based on the digital audio system and has three main systems: • The radio communication system,
The radio communication system enables the cockpit crew to talk air-to-air or air-to-ground via VHF and HF. It is the source of navigation data and ATC transponder function. The Flight and interphone system enables communication among the crew, and to ramp personal when the aircraft is on the ground. The Passenger announcement system enables the cockpit and cabin crews to make live announcements and to play prerecorded announcements, via cabin loudspeakers located throughout the passenger cabin.
System Description The system contains two modular radio cabinets. Within each radio cabinet, radio functions are contained in line replaceable modules. Each major function has its own module with self-contained power supply, RF receivers/ transmitters, signal processing, and all other circuitry necessary for the radio function to operate.
• • • • •
Digital audio bus interface in the MRC Control and display interface to the MCDU radio tuning page Control and display interface to the EDS for radio tuning(MCDU) Control and channel selection via the audio control panel Communication management unit
Figure 1: Components
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Communication system The Digital Audio System consists of three audio control panels, two Network Interface Module (NIM) II in the modular radio cabinets interfacing with the MAU’s through the ASCB-D buses, cockpit loudspeakers, headsets and microphones.
Figure 2: Locations FORWARD E- BAY
MID E- BAY
MRC # 1
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MRC # 2
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General Description The VHF COMM system can transmit and receive voice signals and data in these modes: • Analog voice 8.33 kHz channel (ARINC (Aeronautical Radio Incorporated)-716 type) • Analog voice 25 kHz channel (ARINC-716 type) • Analog MSK (Minimum Shift Key) data mode 0 (ARINC-716 type) • Data link mode A, MSK (ARINC-750 type) • Data link mode 2, D8PSK (Differential 8-Phase Shift Key) (ARINC750 type) The aircraft has three VHF COMM systems: • VHF COMM 1 system • VHF COMM 2 system • VHF COMM 3 system Each VHF COMM system has one VHF COMM module, which is an LRM (Line Replaceable Module). For the VHF COMM 1 system, the module is installed in the MRC (Modular Radio Cabinet) 1, in the forward avionics compartment. For the VHF COMM 2 system, the module is installed in the MRC 2, in the middle avionics compartment. For the VHF COMM 3 system, the module is installed in the MMRC (Mini Modular Radio-Cabinet), in the aft avionics compartment. The DC (Direct Current) ESS (Essential) Bus 1 supplies the power to the VHF COMM 1 system. The DC Bus 2 supplies the power to the VHF COMM 2 system through the SPDA (Secondary Power Distribution Assembly) 2. The DC Bus 2 also supplies the power to the VHF COMM 3 system through the SPDA 2. The VHF COMM system can be tuned by the MCDU (Multifunction Control Display Unit) or by the CCD (Cursor Control Device) and PFD (Primary Flight Display). The MCDU is the primary controller and the CCD and PFD are the secondary controllers. The DAP (Digital Audio Panel) controls the radio selections and audio outputs of the VHF COMM system.
.·
Figure 3: Communication System Schematic
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System Network The radio system network interface is implemented using the NIM module. The NIM is a line replaceable module within the Modular Radio Cabinet and contains the system network bus interface and the digital audio bus interface. The NIM consists of an interface to ASCB that uses the same type of circuitry and software design as the MAU NIC. In addition, the NIM utilizes the standard network Data “mapping” algorithms from the MAU- hosted modules (e.g I/O modules) to interface with the radio bus and the ARINC bus that provides the MRC with the data interface to the Primus Epic system.
MAU Interface The MAU receives the radio data from an ARINC 429 bus at the Generic I/ O Module. The Generic I/O Module contains a standard interface circuit that transfers data to and from a back plane bus. The interface circuit performs functions that include data distribution, data integrity checking, and source identification. The back plane bus is a parallel high capacity general-purpose bus contained in the MAU that transfers all data between the modules and the network interface controller. The network interface controller also contained in the MAU is a dedicated module that interfaces the back plane bus to the external ASCB network. The ASCB network provides the communication between the MAU and the DU`s, where the information is displayed on the PFD’s.
Figure 4: System Architecture
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The Audio Panels Three Audio Control Panels (ACP), one for each crew member and one for the observer provide the audio selection for all audio communications to and from the cockpit. They also provide aural warnings from the MAU and other systems such as TCAS. ACP1 is powered from DC ESS BUS1, and ACP2 from DC ESS BUS2, while ACP3 is powered from the DC BUS 1. The desired channel is selected via different buttons on the ACP. The microphone selection buttons enable the pilot to select any radio to speak. A light illuminates inside the button to indicate that it is pressed. Only one microphone selection button can be selected at a time. The audio selection buttons enable the pilot to select any radio, navigation aid or interphone channel for listening. A light illuminates inside the button to indicate that it is pressed. Any number of selections can be made simultaneously. The telephony buttons on the right - EMER, RAMP, CABN are set up with a hot mic, which provides interphone communication on a telephony basis. A light illuminates inside the button to indicate that it is pressed. Flashing of the light indicates a waiting call from or to the respective station. More then one telephony button can be selected at a time with a maximum of one microphone selection button active.
Figure 5: The Audio Panels
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The Audio Panels (continued) Operation AIRBORNE AUDIO SYSTEM The DAP has the function that follow: • Supplies integrated control of all audio functions in the aircraft. • Receives digital audio data from the NAV and COMM systems through a digital audio bus. • Converts digital data to analog signals that operate the cockpit loudspeakers and headsets. • Supplies an audio output for the DVDR (Digital Voice- Data Recorder) units. The backup mode is set if the DAP power or digital audio bus has a failure. The backup mode connects the microphones directly to the VHF COMM receiver and headphone volume control, releases all transmit buttons, and disarms all other DAP functions and modes. BKUP VOLUME The BKUP Volume switch puts the headset volume in the normal (NORM) or emergency (BKUP) mode. The audio is in NORM mode when the BKUP switch is set to the NORM position. The emergency BKUP mode is set when the switch is released (BKUP- extended position). In the emergency operation mode, power is not necessary to the DAP. The BKUP switch knob adjusts the audio volume to the headset in the BKUP mode of operation. MIC SELECT The MIC select switch gives the operator an alternative between two modes: AUTO or MASK. The headset boom microphone is on when the MIC switch is set to the AUTO position. The mask microphone is on when the MIC button is released (MASK- latched out position). During the emergency operation, the switch can be set to either AUTO or MASK.
LCD The LCD shows audio status data. The audio status data includes volume level, stuck microphone message, and transmit annunciation. MASTER VOLUME CONTROL The master volume control adjusts the loudspeaker volume and sidetone volume. The volume knob is turned clockwise to increase the volume. During the emergencyoperation, the loudspeakers is not available. VHF RADIO MODE The operation to transmit or listen to the audio from the VHF COMM radios are the same for all VHF channels. The operation of the DAP with the VHF1 radio is given below. To hear the audio from the VHF1 radio, the operator pushes the VHF1 AUD button on the DAP. If the SPKR button is selected, the DAP sends the audio to the cockpit loudspeaker. If the HDPH button is selected the DAP sends the audio to the headset. The operator pushes the VHF1 AUD button a second time to stop the audio from the VHF1 radio. To transmit on the VHF1 radio, the operator pushes the VHF 1 MIC button on the DAP. The light on the MIC button goes on. The DAP turns on the audio for that channel (The light on the AUD button comes on). The DAP also adjusts the audio to a preset minimum volume so that the operator will hear any activity on the radio channel. Only one DAP can transmit on the radio while it is in use. The volume display on these other DAPs will show the word BUSY. However, the pilot DAP is set by configuration to override transmissions on the copilot and observer DAPs. The copilot DAP is set by configuration to override transmissions on the observer DAP.
Figure 6: Audio Panels (continued)
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The Audio Panels (continued) When the VHF 1 radio is set to transmit, the volume display shows the 4digit radio code VHF 1 followed by the volume level. The volume level is a number between 0 and 99. The VOL knob adjusts the audio volume while the channel is on and shows in the LCD. The operator pushes the PTT button connected to each DAP to transmit from the headset boom microphone or the mask microphone. While the PTT button is held down, the volume display shows the radio code and the letters TX. If the PTT button is held for longer than two minutes, the radio transmit function ends and the volume display shows the words STK MIC. To stop the microphone inputs to the VHF 1 radio, the operator pushes the VHF 1 button on the DAP. This will also stop the audio from that channel. The lights on the MIC button and AUD button go off. PA MODE The operations to transmit or listen to the audio from the PA system are given below. To hear the audio from the PA system, the operator pushes the PA AUD button on the DAP. If the SPKR button on the DAP is set, the DAP sends the audio to the cockpit loudspeaker. If the HDPH button is set, the DAP sends the audio to the headset. The operator pushes the PA AUD button a second time to stop the audio from the PA system. To speak on the PA system, the operator pushes the PA MIC button on the DAP. The lights on the MIC button and AUD button go on. The DAP turns on the audio for that channel. Only one DAP can transmiton the PA system while it is in use. The volume display on these other DAP will show the word BUSY. EMERGENCY INTERPHONE MODE The flight crew uses the emergency interphone to talk to the flight attendants in an emergency. A flight attendant can also use the emergency interphone to send a signal and speak to the flight crew in an emergency.
CABIN INTERPHONE MODE The flight crew uses the cabin interphone to get communication with the flight attendants. A flight attendant can also use the cabin interphone to call and speak to the flight crew. NAV AUDIO MODE The DAP has audio buttons for the receive- only NAV radios. To listen to the audio from the NAV radio, the flight crew member pushes the NAV audio button. The light in the button comes on. The LCD shows the NAV radio code and the volume level. The VOL knob on the DAP can be used to adjust the volume level. After about 15 seconds, the LCD changes back to show the last active microphone. To stop the audio, the flight crew member pushes the NAV audio button twice, once to energize the channel and the second time to turn it off. ID FILTER MODE The identification (ID) button on the DAP removes the voice signals from the VOR (VHF Omnidirectional Range) or the ADF audio. When the ID filter button is pushed, the light in the button turns on. A filter in the DAP removes all voice signals, the flight crew member pushes the ID filter button again. The light in the button goes out and the audio contains both voice and identification signals.
Figure 7: Audio Panels (continued)
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Optional SELCAL System The digital audio system also includes the selective calling (SELCAL) function. Aircraft implementing the SELCAL function are assigned a unique 4letter SELCAL code. The SELCAL code is set and stored in the AV- 900 configuration database file. The SELCAL function monitors all incoming calls for a SELCAL encoded tone. The function compares the tone sequence of an incoming call to the SELCAL code to find out if it is a selective call. When a call is received, the SELCAL function supplies an aural and visual message to tell the crew that a HF or VHF radio has received a SELCAL transmission. Operation SELCAL OPERATION The SELCAL function has a decoder that lets the ground stations, which have tone encoding equipment , to call an aircraft. The ground stations transmit the unique four-tone code (address) that is given to the aircraft. These encoded audio tones only operate with a SELCAL function set to that combination of tones. The SELCAL code is a BCD (Binary Coded Decimal) four-letter code. This combination code is composed of four audio tones that must be transmitted by the ground station to alert the crew. Each transmitted code from the ground station has two consecutive pulses. Each pulse contains two tones (two letters). Each code letter is an audio tone in BCD format. Each aircraft is given an individual SELCAL code by the ARINC (Aeronautical Radio Incorporated) registration. When the aircraft receives a selective calling, the lights in the SELCAL button, MIC and AUD control buttons of the applicable HF or VHF COMM radio start to flash on the DAPs, and an aural message is heard in the cockpit. By pushing the MIC button of the applicable HF or VHF COMM radio, the cockpit crew is able to speak into the radio and answer the call. The lights in the related MIC and AUD buttons and SELCAL button becomes stable on. By pushing the SELCAL button, the communication stops. SATCOM Operation (optional) The operation of SATCOM selection through the audio panel is the same as radio (VHF and HF) selection except for the following differences:
When a call is received (call dixcrete from the SATCOM), the annunciator on the SAT microphone button will flash. An aural annunciation will be provided by the audio system to the HDPH and SPKR channels. When the SAT microphone button is selected, both SAT annunciators turn on, and a discrete is output to the SATCOM to answer the call. The headset or mask Mic is then “hot Mic”, and use of PTT is not required, unless the hand Mic is used. The pilot may talk and listen as desired.
Figure 8: Digital Audio Panel Control/ Indicators
1. MIC BUTTON (6x)
2. AUD BUTTON (13x)
3. MKR BUTTON
4. ID BUTTON
MIC
VHF1
VHF2
VHF3
HF
PA
SAT
AUD
5. EMER BUTTON NAV1
NAV2
NAV3
DME1
DME2
MKR
ADF1
ADF2
ID
EMER
15. SELCAL BUTTON 14. MIC SWITCH
SELCAL
RAMP
SPKR
INPH
HDPH
6. RAMP BUTTON
CAB
7. CAB BUTTON 13. BKUP SWITCH
BKUP
VOL
MIC
VHF1: 47 NORM
BKUP
AUTO
8. VOL KNOB
MASK
9. HDPH BUTTON 12. LIQUID-CRYSTAL DISPLAY
DIGITAL AUDIO PANEL
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10. INPH BUTTON 11. SPKR BUTTON
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Figure 9: Digital Audio Panel Control/ Indicators REF.
AUD (audio-receive) Button (VHF 1, VHF 2, VHF 3, HF, SAT, PA, NAV 1, NAV 2, NAV 3, ADF 1, ADF 2, DME 1, DME 2) MKR (marker) Button
ID (identification) Button
EMER (emergency) Button
RAMP Button
CAB (cabin) Button
8
VOL (volume) Knob
9
HDPH Button
POSITION/INDICATION
MODE
selected
Connects the MIC and arms the audio function of that channel. The button has a light that comes on when the MIC is on. For channels with telephone features, the light flashes for incoming calls or to indicate hold status.
deselected
Disconnects the microphone of the channel.
selected
Turns on the audio function of the related channel. The button has a light that comes on when the audio is on.
deselected
Turns off the audio of the channel.
selected
Turns on the marker audio signal.
deselected
Turns off the marker audio. The button has a light that comes on when the audio is on.
selected
Removes the voice audio from the receivedVOR and ADF audio. The button has a light that comes on when the audio is on.
deselected
Removes the voice filter. The Id audio and voice signals are both heard.
selected
Used to call the cabin attendants in an emergency. Connects the MIC and arms the audio input. The button has a light that comes on when the emergency interphone channel is on.
deselected
Turns off the emergency interphone audio.
selected
Used to get communication with the maintenance crew. Connects the MIC and arms the audio input. The button has a light that comes on when the ramp interphone channel is on.
deselected
Turns off the ramp interphone audio.
select
Used to get communication with the flight attendants. Connects the MIC and arms the audio input. The button has a light that comes on when the cabin interphone channel is on.
deselected
Turns off the cabin interphone audio.
clockwise to increase volume
Adjusts the volume of the last active audio channel. The volume level and active channel show in the LCD.
selected
Connects the audio signal into the headset.
deselected
Disconnects the audio signal from the headset.
Figure 9: Digital Audio Panel Control/ Indicators (continued)
REF. 10
11
CONTROL/INDICATOR INPH (interphone) Button
SPKR Button
POSITION/INDICATION
MODE
selected
Lets the pilot, copilot, and observer speak on a cockpit interphone channel.
deselected
Turns off the cockpit interphone channel.
selected
Connects the pilot or copilot audio signals into the loudspeaker.
deselected
Disconnects the audio signal for the loudspeaker. This LCD shows the code for the active COMM channel and one of the following:
12
13
14
15
LCD
BKUP (normal/backup) Switch
(none)
last update: Dec06
Volume level - 0 to 99 Status - BUSY, TX (transmit), CON (connected), STK MIC, NO SOUND
NORM (latched-in)
Sets the DAP for normal operation.
BKUP (latched-out)
Sets the DAP for the emergency radio operation mode. Connects the MIC directly to the VHF COMM transmitter. At the same time, the headset connects to the VHF NAV/ VHF COM receiver and the HDPH volume control. All the MIC (transmit) buttons will be released. All the other DAP functions and modes will be turned off. In the BKUP mode, the BKUP knob controls the audio volume. The light in the BKUP switch goes on when the switch is in the BKUP mode.
AUTO (latched-in)
Energizes the boom MIC and headset. When the pilot puts on the oxygen mask and starts the air flow, the oxygen mask microphone turns on and the headset boom microphone goes off. If the auto mode has a malfunction, the oxygen mask microphone can be set manually when you push the MIC switch to the MASK position.
MASK (latched-out)
Energizes the oxygen mask MIC. The light in the MIC switch goes on when the switch is in the MASK mode.
selected
Resets the ground-to-air calling system indications. This function supplies an aural warning and an EICAS (Engine Indicating and Crew Alerting System) message when a preset 4-letter code is received by one of the COMM radios. The light in the SELCAL button flashes when an incoming call is received. The pilot and copilot do not have to continuously monitor the communication frequencies.
MIC (auto/mask) Switch
SELCAL Button
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RAMP Interphone Operation The operation of RAMP selection through the audio panel is defined as follows: • Pressing a button on the pilot’s left console or on the copilot’s right console sounds a horn in the nose wheel well to summon ramp personnel. When ramp is selected on the audio panel, and hot microphone is enabled, the pilot microphone signal is routed to the PACIC. There are provisions for three ramp service (maintenance) interphone connections. • If the pilot desires, he can select another radio, at this time the ramp annunciator on the ramp select button remains lit. To return to the ramp function, the ramp select button is pushed. • For an incoming call, the ramp annunciator flashes. To pick-up the call push the ramp select button, at this time the annunciator is steady on and the display shows the characters denoting RAMP and a number between 0 and 99 denoting the volume.
Figure 10: Ramp Interphone Operation
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RAMP CALL HORN The ramp call horn is installed in the nose-landing-gear wheelwell. It sounds when the pushbutton on the cockpit left or right console is pushed.
Figure 11: Ramp Interphone COCKPIT LEFT CONSOLE HORN PUSHBUTTON
Radio Tuning Normal Operation Normal Radio Tuning is conducted through pages on the MCDU. There is a dedicated button for RADIO on the MCDU and when selected it displays the VHF Comm page, from this page there is a Radio Menu Soft Key which displays a menu to select any of the other Radios. When the RADIO button is selected on the MCDU two pages of radio information are displayed and a detailed page for each radio is also available.
Figure 12: Radio Tuning, Normal Operation
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Radio Tuning Abnormal Operation Backup Tuning for the Radios is conducted through the Cursor Control Device and is displayed on the Primary Flight Display. If there is a NIM1 failure, a back-up ARINC 429 bus may be used to tune VHF COM 1,NAV1 and XPDR 1 from MCDU 2.The digital audio buses are redundant and contain identical information so the failure of a single NIM or bus will not interrupt the pilot/co-pilot audio signals. Loss of communication on the ASCB bus is denoted by dashes in the MCDU radio tune window. A radio NIM failure is enunciated as a CAS message; NAVCOM 1(2) FAIL.
Figure 13: Radio Tuning Abnormal Operation
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23-11 High Frequency HF Communication System (ATA 23-11-00) / optional The EMBRAER 170/190 is equipped with a single KHF-1050 HF system, capable of long range air-to-air and air-to-ground voice communications. The system consists of the following major LRUs: • • • •
Tranceiver Unit (KRX-1053) Power Amplifier (KPA-1052) Antenna Coupler (KAC-1052) HF Shunt Antenna (integral part in the leading edge of the vertical fin)
The HF system is integrated with the aircraft‘s avionics system, allowing frequency tuning & control via the two flight deck MCDUs, and audio control via the audio panels. The Transceiver and Power Amplifier units are installed in the aft e-bay rack. The Antenna Coupler is also installed in the aft e-bay ceiling area, near the antenna, to maximize transmission efficiency. The aft e-bay is pressurized and temperature controlled.
Figure 1: HF Component Location (Aft e-Bay)
C B
D
A
A
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HF TRANSCEIVER
HF POWER AMPLIFIER
HF ANTENNA COUPLER
B
C
D
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HF LRU/ Component Description Transceiver Unit (KRX- 1053) The Transceiver unit contains an exciter, which produces a very low power signal, which is fed to the Power Amplifier. The Receiver/ Exciter handles system control communication, mode control, channelization, and controls transmission power. During transmission a TX annunciation is displayed beside the active HF frequency. Power Amplifier (KPA- 1052) The Power Amplifier amplifies the signal from the Receiver/ Exciter into a 200 watt Peak Envelope Power (PEP) transmitted signal in single sideband operation or 50 watts in Amplitude Modulation (AM) operation. After filtering, the amplified signal is fed to the Antenna Coupler. In the receive mode, the signal is passed to the Receiver/ Exciter triple conversion Intermediate Frequency (IF) Amplifiers and then digitized for higher fidelity demodulation. Antenna Coupler (KAC- 1052) The Antenna Coupler is mounted near the antenna to maximize transmission efficiency. A microprocessor in the Antenna Coupler controls the Power Amplifier and exchanges information with the MCDU via the Receiver/ Exciter. The Antenna Coupler matches the impedance of the antenna to the 50 Ohm output of the transmitter. The coupler will seek a SWR of 1:1, but anything lower than 2.5:1 can be considered tuned. To facilitate operation at altitude the antenna coupler is pressurized to prevent internal arcing. Should pressurization drop below nominal limits, the power amplifier will reduce transmission power to a safe level. Should pressurization fail completely the HF system will disable transmission until the aircraft descends to an altitude with sufficient air pressure to prevent arcing. Since HF transmission requires the tuning of the antenna coupler to the selected frequency, a tuning cycle is triggered within the antenna coupler whenever a new frequency is selected on the MCDU and the PTT pressed. A TUNING message is annunciated during the tune cycle on the MCDU HF detail page 1 of 2, in inverse video. Additionally, an aural indication of the tune cycle is annunciated by a 1000 Hz tone in the headset, until tune is complete.
New frequency tuning should occur within 30 seconds, with rare cases taking longer. Retuning to a previously tuned frequency should occur within 1 second. Should the system fail to tune, this will be annunciated on the same page in inverse video and a scratchpad message will be generated, “TUNE FAIL”. This annunciation will remain until either a new tune cycle is initiated by re- keying the mike, or a new frequency is selected to tune. HF Shunt Antenna The Embraer 170/ 190 HF Shunt antenna is an integral part of the leading edge of the vertical fin. It is comprised of a floating (isolated) folded metalic foil sheet that is attached at its lower end to the antenna coupler feed thru adapter via a metal strip, while its top end is grounded at a convenient point of the aircraft. The HF antenna dimensions and design are based primarily on previous shunt antenna models (from other Embraer aircraft programs), but enhanced for superior performance with recommendations provided by Honeywell‘s HF engineering team.
Figure 2: HF System Block Diagram & Architecture
MCDU 1
ARINC 429
MAU 1
MAU 3
ARINC 429
MAU 2
CONTROL I/O
GENERIC I/O
CONTROL I/O
NIC
NIC
NIC
MCDU 2
OBSERVER DAP MRC 2
MRC 1 AUDIO BUS
SELCAL AUDIO
AUDIO
NIC
MIC BUS
MIC
NIC
HF CONTROLLER 2
COPILOT DAP
TUNING MODE
PILOT DAP
HF PTT
HF TUNING/MODE STATUS
HF CONTROLLER 1
ASCB
HF TRANSCEIVER
HF ANTENNA
COUPLER BUS
CONTROL/ STATUS
HF AUDIO HF ANTENNA COUPLER
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HF AUDIO
PA SERIAL DATA CONTROL/STATUS
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HF POWER AMPLIFIER
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HF Antenna The HF antenna is a shunt-type antenna installed as part of the leading edge of the vertical stabilizer.
Figure 3: HF ANTENNA LOCATION
A
HF ANTENNA
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Modes & Control The two flight deck MCDUs provide the primary man- machine interface to the HF system. It enables the pilots to access, tune and control the system‘s operational modes, while audio (volume) control is performed via the flight deck AV- 900 audio panels. Voice transmission can be made in conventional upper and lower sideband. These are denoted on the MCDU as UV and LV for Upper Voice and Lower Voice. Amplitude Modulated communication is supported through the AM mode, and RC denotes reduced carrier power single sideband for frequency tracked receivers. With an external modem, data is supported in either in upper or lower sidebands, denoted UD or LD, however this mode of operation is not employed on the Embraer 170/ 190. Voice operation can be either Simplex, Duplex or ITU channelizations. These operational modes are designed as follows on the MCDU screen: • SMPL -simplex transmission and reception on the same frequency. • SPLT -duplex operation in which transmission and reception frequencies differ. • ITU -duplex operation with frequencies in accord with ITU channel designations. • EMERG -single button access to six pre- stored emergency frequencies. There are four methods available for noise reduction, these are listed below: • SQL -signal strength squelch, for a low noise and high signal environment. • SQH -audio noise squelch, for a high a noise environment. • SBL -audio frequency syllabic content squelch, for a low signal environment. • SBH -audio frequency syllabic content squelch, for a high signal environment.
Figure 4: NORMAL OPERATION
Figure 4: Radio Page 1 of 2
TX
Figure 5: Radio Page 2 of 2
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23-24 Communications Management Function Introduction The CMF (Communications Management Function) is an ACARS (Airborne Communication Addressing and Reporting System) network compatible router through which the character-oriented data messages are communicated between the ground datelined service providers and aircraft systems. The CMF supplies a digital interface with the flight crew for the operation and management of the COMM (Communications) radio systems.
Figure 1: CMF Overview
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General Description The CMF is a software function that is hosted on the processor modules in the MAU (Modular Avionics Unit)s. It hosts the functional software and communications protocols that let the CMF send messages between the airborne and ground-based subnetwork systems, and operates as an end-system host. As a host, the CMF supplies the datelined application functions and router tasks for the ATS (Air Traffic Services) and AOC (Airline Operational Communications). The other datelined end-systems in the aircraft are the FMS (Flight Management System) and CMC (Central Maintenance Computer). The CMF stores the ATS and AOC messages so they can be set by the flight crew and show on the MCDU (Multifunction Control Display Unit) in the cockpit. The flight crew can use the CMF pages on the MCDU to start requests for the data and send messages to the ground systems. The CMF communicates to the systems external to the MAU through the ARINC (Aeronautical Radio Incorporated)-429 buses, through the control I/O (Input/Output) module in the MAU, or LAN (Local Area Network). The CMF communicates to the systems internal to the MAU through the ASCB (Avionics Standard-Communication Bus). The CMF communicates with the other elements of the system through the virtual backplane (PCI bus). The CMF has the capabilities that follow: • Control the data communications over the air/ground subnetworks. • Report the communications status and availability of each air/ ground subnetwork. • Host ARINC-623 character-oriented ATS applications such as the clearances and ATIS (Automatic-Terminal-Information Service)s. • Supply the capability to let the airline/aircraft OEM (Original Equipment Manufacturer) customize the AOC applications through the creation of a downloadable AMI (Airline Modifiable Information)
database. (This supplies control of the MCDU display screen formats, print definitions, and message contents and formats). • Connect to the LAN printer so the messages can be printed. • Report the faults and events to the CMC. • Record the faults and events in NVM (Non-Volatile Memory). • Supply the alert indications to the CAS (Crew Alerting System). The CMF has the component that follows: • Printer
Figure 2: CMF Block Diagram
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CMF Datalink Application- Air Traffic Service (ATS) The CMF operates with the ATS applications, which are the characteroriented applications that transmit the data on the ACARS network. The CMF supplies the display and message procedure associated with the ARINC-623 ATS applications. The CMF supports the ATS applications that follow: • The ATIS application is used to supply the terminal data reports to the flight crew. These messages are to request an uplink report and to show that data to the flight crew. The ATIS request downlink message is supplied in response to the crews action. This downlink message requests an ATIS uplink report for a specific airport or for enroute data. The ATIS report uplink message is a response to an ATIS request downlink. • The TWIP (Terminal Weather Information for Pilots) application is used to supply the weather data reports to the flight crew. These messages are to request an uplink report and to show that data to the flight crew. There are two TWIP messages , one uplink and one downlink. The TWIP request downlink message is supplied in response to the crews action. This downlink message requests a TWIP uplink report for a specific airport. The TWIP report uplink message is a response to a request downlink. • The oceanic clearance application is used to request an oceanic clearance that uses the ground system to supply that clearance. These messages are to replace and supplement the voice communications that would normally be used for this type of clearance. There are three oceanic clearance messages, one uplink and two downlinks. The oceanic clearance request downlink message is supplied in response to the crews action. This message requests an oceanic clearance uplink. The oceanic clearance uplink message is a response to an oceanic clearance request downlink. The oceanic clearance readback downlink message is completed in response
to the crews action. This message is used as an acknowledgement of the oceanic clearance uplink.
The CMF supplies the display and message procedure associated with the ARINC-623 ATS applications. The CMF supports the ATS applications that follow: •The expected taxi clearance request downlink message is supplied in response to the crews action. This message requests a taxi clearance uplink. The expected taxi clearance uplink message is a response to an expected taxi request downlink. • The pushback clearance application is used to request a pushback clearance that uses the ground system to supply that clearance. These messages are to replace and supplement the voice communications that would normally be used for this type of clearance. There are two pushback clearance messages, one uplink and one downlink. The pushback clearance request downlink message is supplied in response to the crews action. This message requests a departure clearance uplink. The pushback clearance uplink message is a response to a pushback clearance request downlink.
Figure 3: CMF Functions
1.ACARS - Aircraft Communication and Addressing System CMF - Communicates with Ground Stations on the ACARS network.
CMF - Operates with ATS (Air Traffic Service) Applications that Include: ATIS - Airport info (QNH,QFE,Dew point etc). TWIP - Terminal Weather Info for Pilots. Taxi Clearance. Push Back Clearance.
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Operation with other Systems The CMF operates with the systems that follow: • ACARS — The CMF uses the applicable communications protocols (i.e., ARINC-618 air/ground protocols and ARINC-750 VDR (VHF Data Radio) protocols) to process the data messages it receives through the ACARS datelined. If the message text identifies the CMF as the system to receive the message, the CMF processes the message. If the message text identifies a system other than the CMF as the system to receive the message, the CMF removes the message text from the message (or messages from a multiblock message). The CMF then adds the control/ accountability header to the start of the message and sends the message to the applicable system. • Visual warning function— The visual warning function has a CAS that monitors the CMF function for alert indications. When the new data messages have been received and are ready to be read, the CMF sends an alert indication to the CAS to notify the pilot or copilot. The CAS operates with the EICAS (Engine Indicating and Crew Alerting System) display to show the alert indications. When a system failure occurs, the EICAS display in the central display system shows the CAS messages. • Central Maintenance System — System faults are transmitted to the CMS where the CMC records them. A portable computer can be connected to the CMC processor module through the LAN to read the system fault data. CMF INTERFACES - FLIGHT MANAGEMENT SYSTEM (FMS) The CMF operates as a router for the FMS messages to and from the ground systems. The CMF receives messages from the subnetworks through the applicable radio and MAU control I/O module. If the message is for the FMS (end system), the CMF removes the message
text, adds a header, and sends the message to the FMS. The FMS interfaces with the CMF through the ASCB. When a message is sent from the FMS to the ground system, the path is the same, but the direction of the message is reversed. The FMS sends the messages that follow to the CMF for transmission to the ground systems: • Flight plan requests • Flight plan reports • Position reports • Requests for winds aloft • Response messages • Rejection messages The FMS receives these messages from the ground systems through the CMF router: • Flight plans • Winds aloft • Position reports at waypoints The definition of the messages communicated to the FMS is specified by the FMS, not the CMF. CMF INTERFACES - CENTRAL MAINTENANCE COMPUTER (CMC) The CMF operates as a router for the CMC messages to and from the ground systems. The CMC is part of the CMS. The main function of the CMC is to supply the computer resources for the other systems on board the aircraft.
Figure 4: CMF Functions
2. CMS - Central Maintenance System CMF - The Central Maintenance system uses the CMF to send fault Messages and Reports to the ground Stations.
3. FMS - Flight Management System CMF - The FMS uses the CMF to send and receive: Flight Plan Requests. Position Reports. Winds aloft.
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Access to the CMF pages The MCDU is the primary flight crew interface that supplies the display and control for the CMF. To access the CMF pages, push the datelined function key (DLK) on the MCDU. If it is the first time that the CMF pages have been accessed since power-up, the CARS MAIN MENU page shows. If it is not the first access since power-up, and there are new messages, the NEW MESSAGES page shows. If it is not the first access since power-up, and there are no new messages, the last CMF page accessed shows. CARS MAIN MENU The CARS MAIN MENU is the root CMF page. It shows on the MCDU when the DLK function key is pushed for the first time after power-up. Any of the AOC pages that show in this menu can be selected. The AOC pages show the contents of the AMI database. These pages are customized by the airline when the AMI database is loaded with the GBST. The LSK to the left or right of each line in the CARS MAIN MENU selects that page as the next one to be shown on the MCDU. In this way, the CMF pages are navigated from the high-level AOC menu to the pages on the low-level menus. The most commonly accessed AOC pages in the AMI database are available from the CARS MAIN MENU and are listed as follows:
• • • • • • • • • • • •
PRE FLT - pre-flight menu IN FLT - in-flight menu POST FLT - post-flight menu FREE TEXT - free text display FLT TIMES - flight times display SYS MENU - system menu access NEW MSGS - new messages display MSGS SENT - shows messages that have been sent MSGS RCVD - messages received log VOX CONTACT - VHF voice contact request display STATUS - status menu display. ATS MENU - ATS menu display
The ACARS MAIN MENU also has line prompts to request the SYSTEM and ATS menus.
Figure 5: CMF MCDU Operation
MA I N ME NU
ACARS PRE
FLT
N EW M S G S
FLT
IN
POS T
FLT
FREE
TEXT
FLT
MSGS
SENT
MSGS
RCVD
VOX
T I ME S
S T A TUS
S Y S ME NU
PERF
NAV
CON T A C T
PREV
A T S ME NU
FPL
PROG
DIR
BRT
CB
A
DIM MENU
DATA LINK FUNCTION KEY-TO SHOW CMF PAGES
DLK
NEXT
TRS
RADIO
A
B
C
D
E
F
G
H
I
J
K
L
1
2
3
+/ -
M
N
O
P
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/
S
T
U
V
W
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8
9
X
Y
Z
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CLR
SP
0
MCDU
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Printer The CMF includes a medium-speed dot matrix printer that uses thermally sensitive paper. The printer has data and control electronics, a power supply, and a paper supply/control mechanism. The printer is mounted in the cockpit on the copilot side of the control pedestal. During maintenance activities, the printer is used to print the maintenance data and other diagnostic data from the CMS (Central Maintenance System). All communications between the printer and other systems are directed by the CMC. The MCDU menus supply the option to print the CMF display messages on the printer, over the LAN. The printer operates from 28 VDC (Volt Direct Current) power supplied by the DC (Direct Current) BUS 1. The printer front panel has a power on/off switch and control indicators for operator control, test functions, and status display. When the door on the front of the printer is open, access to the paper supply for the unit is available. The printer holds a 125 ft roll of paper. The low-paper indicator on the front panel comes on when about ten ft of paper is left on the roll. The last 6 ft of paper on the roll has a colored warning stripe, which does not affect the legibility of the printed data. If the paper runs out, the fault and low-power indicators come on and the printer does not print until the paper is added. The detection of a paper-out condition stops further printing and the data in the printer electronics buffer is retained until the paper is replaced.
Figure 6: Printer
PRINTER - FOR PRINTING CMF (ACARS) MESSAGES. - FOR PRINTING CMC MAINT DATA. Issue: June06 Revision: 00
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23-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 50-00 Crewmember ¦ Interphone ¦ Systems
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: ORIGINAL ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 12/16/2003 ¦ 23-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 23 COMMUNICATIONS ¦ ¦ ¦
¦ 11-00 High Frequency ¦ *** (HF) ¦ Communication ¦ System
D ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ Any in excess of those required by ¦ FAR may be inoperative. ¦ ¦
¦ ¦ ¦ ¦
¦ 12-00 Very High ¦ Frequency (VHF) ¦ Communication ¦ System ¦
D ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ Any in excess those required by FAR ¦ may be inoperative provided: ¦ a) VHF 1 is operative, and ¦ b) Procedures do not require ¦ its use.
¦ ¦ ¦ ¦ ¦
¦ 15-00 Satellite ¦ *** Communication ¦ System (SATCOM)
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
¦ ¦
D ¦ ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ ¦
¦ 21-00 Selective Call ¦ *** System (SELCAL) ¦
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided flight ¦ crew continuously monitors ¦ appropriate radio frequencies.
¦ ¦ ¦
¦ ¦
D ¦ ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ 24-00 Communication ¦ *** Management ¦ Function (CMF)
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ ¦
D ¦ ¦
¦ 24-01 Printer ¦ *** ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
1) Flight Deck to B ¦ Cabin, Cabin ¦ to Flight Deck ¦ Functions ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Flight deck to cabin and ¦ cabin to flight deck ¦ interphone functions operate ¦ normally on at least fifty ¦ percent of the cabin ¦ handsets, and ¦ b) Alternate communication ¦ procedures between the ¦ affected flight attendant ¦ stations are established and ¦ used.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ NOTE: ¦ ¦
¦ ¦
¦ 51-01 Cockpit Alerting ¦ System ¦ (Chime/Light)
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
¦ ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided the ¦ flight deck chime operates ¦ normally.
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ ¦
¦ ¦
¦ ¦
¦ NOTE: ¦
The flight deck chime must always be operative.
¦ ¦
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
D ¦ ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ ¦
¦ 51-02 Cockpit Speakers ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ May be inoperative provided: ¦ a) Procedures do not require ¦ its use, and ¦ b) Headsets are installed and ¦ operate normally.
¦ ¦ ¦ ¦ ¦
1) Flight Deck Call Lights
B ¦ 3 ¦ ¦ ¦ ¦ C ¦ 2 ¦ ¦ ¦ ¦
Any station function(s) that operate normally may be used.
44-00 Crew communication Table 1: ATTENDANT CALL INDICATORS- GENERAL DESCRIPTION
Introduction
COLOR
DESCRIPTIPON
ORANGE
When a PAX presses the ATDT call SW located inside the lavatory.
The cabin is equipped with two communication stations: one at the fwd and one at the aft left hand side cabin crew seat.
BLUE
When a PAX in the cabin area presses the ATDT call SW on the PSU located above the seat.
The communication stations consist of an interphone handset with an ear piece, electronic microphone and a push-to-talk button. The handset is fixed in its cradle by a magnetic latch.
AMBER
The Passenger Address and Cabin Interphone System provides the capability for communication between the cabin crew, pilots and passengers.
The cradle is provided with four control switches: PA; ATTND; PILOT and EMER PILOT. Attendant call lights located on the fwd and aft main ceiling panel areas provide visual indication to the cabin crew when there is a call from the pilots or passengers.
When the pilot does not want to be disturbed. This light is controlled by a SW installed on the overhead panel in the cockpit, and is designated to illuminate the sterile light.
RED
When the pilot makes an emergency call to the flight ATDT from the cockpit.
GREEN
When the pilot calls the flight ATDT from the cockpit.
FORWARD/ AFT ATTENDANT LIGHT INDICATOR PANEL The ATDT light indicator panel is composed of five different colors, each one with its own meaning and are all assembled on the ATDT light indicator panel on the ceiling panel. There are indicators installed on the FWD and AFT ceiling panels to indicate ATDTs that a call was originated. The ATDT call lights provide a visual indication to ATDT when there is a call from the flight crew or passengers. Each type of call there is just one color, that could be orange, blue, amber, red and green.
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Figure 1: General
TEMPERATURE SETTING
LOW
TEMPERATURE SETTING
OFF
OFF
C
HIGH
D O O R Z O N E T E M P E R AT U R E
H
ON
ON
C E IL IN G
BRIGHT DIM
S ID E WAL L
BRIGHT DIM
TEST
ON
BRIGHT DIM
FWD GALLE Y AREA
PANEL LIGHTS
BRIGHT DIM
BRIGHT DIM
TEST
C OU RTESY LIG H T
RESET
EVAC H O R N
OFF
BRIGHT DIM
ENABLED
ON / ARMED
GALLEY MASTER
OFF
A FT E N TR AN C E
TEST
PANEL LIGHTS
BRIGHT DIM
TEST
C O U R TES Y LI G H T
EMERG ENCY LIGHT
RESET
OFF
E VA C H O R N
ON
AUTO
WA STE SYS TEM
ON
AUTO
FWD
LAVATO RY SMOK E TEST
FWD
H
ON
S ID E WA L L
OFF
FW D E N TR AN C E
EMERGENCY LIGHT
ON / ARMED
C
CABIN LIGHTING
GALLEY MASTER
ON
C A B I N T E M P E R ATU R E
ENABLED
C AB IN TE M P ER AT U R E
CABIN LIGHTING
ON
C E IL IN G
HIGH
D O O R Z O N E T E M P E R AT U R E
LOW
ON
AFT
PSU
TEST
TANK FULL
AFT
SERVICE TANK
FAULT
LAVATOR Y FA U LT
RESET
ATT N D C AL L
WAT ER S YSTE M WATE R QU AN TITY
RESET
FAULT
ATTND CALL
1/4
1/2
3/4
RESET
PA
PA
ATTND PILOT EMER PILOT
PSU
ATDT EMER
STER
ATTND PILOT EMER PILOT
PA
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Passenger address system The Passenger Address (PA) system allows announcements to be made to passengers, either by pilots on the flight deck or the cabin crew members at their stations. The announcements are heard through speakers located in the cabin and in the lavatories. The use of the PA system is prioritized. The announcements from the flight deck have first priority and override all others. The announcements from the cabin override the pre-recorded announcements which override the music system. Announcements in the cabin are performed by lifting the handset from the cradle and pressing and holding the PA button.
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Figure 2: PA system
TND PILOT EMER PILOT
Press and hold for cabin announcements
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Service Interphone system The service interphone system provides intercommunication between the pilots and the cabin crew. To communicate with other cabin crew members the ATTND button has to be pressed. To gain their attention an aural signal is heard and an amber light at the cabin crew panel illuminates. When Pilot is selected, a single CALL chime is annunciated to the flight deck. The same chime is heard in the cabin when the pilots select the option CAB for a regular cabin call. In addition to the aural signal, a green light in the cabin crew panel illuminates.
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Figure 3: Service interphone system
PSU
LAV
ATDT EMER
STER
MER ILOT
+ 1 chime
MIC
VHF1
VHF2
VHF3
A 1 AV
NAV A 2
A 3 AV
DME1
ME2
KR
HF
SAT
VOL
LCAL BKUP VOL
DF1
DF2
EME
ID
CAB
SPKR
INPH
HDPH VOL
MIC
VHF1: 47 BKUP
MAS
Honeywell
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ATDT EMER
STER
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In case of an emergency To announce an emergency, the cabin crew members have to press the PILOT emergency button. This causes a triple chime on the flight deck. In the same manner the cabin will hear a triple chime when the option "emergency" is selected by the pilots on the flight deck. A red light illuminates in the cabin crew panel.
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Figure 4: In case of an emergency
PSU
LAV
ATDT EMER
STER
+ 3 chimes
PA MIC
VHF1
VHF2
VHF3
HF
SAT
A 1 AV
NAV A 2
NAV A 3
DF1
DME1
ME2
KR
ATTND PILOT EMER PILOT
PA
VOL
SELCAL BKUP VOL
DF2
EMER
ID
CABN
SPKR
INPH
RAMP
HDPH VOL
MIC
VHF1: 47 BKUP
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44-11 Passenger Address and Cabin Interphone System Introduction The PACIS (Passenger Address and Cabin Interphone System) controls the communication between the cockpit crew and passenger cabin, cockpit crew and flight attendants and flight attendants and passenger cabin. The PACIS also supplies the logic for generation of aural and/ or visual annunciators for the cockpit crew, flight attendants and passengers.
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Figure 1: PACIC System architecture
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General Description The primary component of the PACIS is the PACIC (Passenger Address and Cabin Interphone Controller), which interfaces with these systems: • Airborne audio system • Ramp interphone • Modular avionics unit • Modular radio cabinet system • Cabin loudspeakers • Cabin interphone • Warning signs • Attendant call indicators • Ambient music system • Prerecorded announcement
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Figure 2: PACIC System Schematic B
A
MAX MIN ON TONE
A PASSENGER ADDRESS CTL PNL (SDS 44-11) (MPP 44-11-02)
Components PASSENGER ADDRESS AND CABIN INTERPHONE CONTROLLER The PACIC does the functions of a cabin interphone controller and passenger address amplifier. Each of these parts operates independently and has independent supplies. It is installed in the forward avionics compartment. The functions of the PACIC are: • To increase the audio level of all signals addressed to the passenger cabin. • To supply the logic for generation of aural and/or visual annunciators for the cockpit crew, flight attendants and passengers. • To control the communication between the cockpit crew and passenger cabin, cockpit crew and flight attendants, and flight attendants and passenger cabin. Training Information Points CAUTION: MAKE SURE THAT THE PONTENTIOMETER FOR OUTPUT GAIN ADJUSTMENT, ON THE FRONT FACE OF THE PACIC, IS IN THE CERT POSITION. IF YOU CHANGE THIS POSITION THE PACIS WILL NOT OPERATE CORRECTLY. When the aircraft is on ground (WOW (Weight-on-Wheels) = true) and the engines are off, the PACIC decreases the audio output to the cabin loudspeakers by 6 dB.
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Figure 3: PACIC. Forward E-Bay
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CABIN LOUDSPEAKERS Introduction The cabin loudspeakers function is to supply the audio amplified by the PACIC (Passenger Address and Cabin Interphone Controller) to the passenger cabin. General Description The cabin loudspeakers are installed in such a manner to assure that all the crew and passengers can hear the addressed sound, independently of their locations in the cabin. Components PASSENGER LOUDSPEAKERS There is one passenger loudspeaker installed on each PSU (Passenger Service Unit). LAVATORY LOUDSPEAKERS The lavatory loudspeakers are installed on the lavatory ceiling. There is one loudspeaker installed in each lavatory. CEILING PANEL LOUDSPEAKERS The ceiling panel loudspeakers are installed at the forward and aft passenger entrance areas. There are two loudspeakers installed on the forward ceiling panel and two on the aft ceiling panel. Operation The cabin loudspeakers operate whenever the PACIC supplies an amplified audio signal or a chime. Training Information Points The lavatory loudspeakers only operate if the lavatory door is closed and latched.
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Figure 4: Cabin loudspeakers
A
G
C
L
K
J
B M D
F
H
E L
C
H
E
B
D
A F WD P AX E NTR A NC E C E IL I NG PA NE L
F WD GAL L E Y A R EA C E IL I NG PA NE L
F WD
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 3)
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 3)
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 2 )
G
H
L AV A T OR Y
1S T RH PS U
3RD RH PS U
5TH RH PS U
13 TH RH PS U
15 TH RH PS U
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
F
F
F
F
F
F
F
F
FL
1
7TH RH PS U
11T H RH PS U
9TH RH PS U
17 TH RH PS U
2 ND RH PS U
4TH RH PS U
6TH RH PS U
12 TH RH PS U
14 TH RH PS U
16 TH RH PS U
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
F
F
F
F
F
F
F
F
F
F OR WA R D A V IONIC S C OM P T
8TH RH PS U
10 TH RH PS U
18 TH RH PS U
A T T #1
P A CI C (S S M 4 4 11- 80 ) A T T #2
F
F
F
F
F
F
F
F
F
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
4TH L H PS U
6TH L H PS U
8TH L H PS U
10 TH L H PS U
12 TH L H PS U
14 TH L H PS U
16 TH L H PS U
18 TH L H PS U
2 ND L H PS U
K
J
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 3)
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 3)
A F T P AX ENT RA NC E C E IL I NG PA NE L
Issue: June06 Revision: 00
A F T R H MAI N C E IL I NG PA NE L
M
1
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 2 ) AF T
L AV A T OR Y
F
F
F
F
F
F
F
F
F
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
L OUDS PE A KE R (S DS 44 - 12 ) (MPP 44 - 12 - 0 1 )
3RD L H PS U
5TH L H PS U
7TH L H PS U
9TH L H PS U
11T H L H PS U
13 TH L H PS U
15 TH L H PS U
17 TH L H PS U
1S T L H PS U
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170/190 MAINTENANCE TRAINING MANUAL
PACIC System FLIGHT-ATTENDANT HANDSET/CRADLE There are flight-attendant handset/cradles in the FWD (Forward) and aft work areas, and they can be installed on the airstair wardrobe or lavatory. On the handset cradles there are four buttons with the related LED (Light-Emitting Diode). Each button has a different function, as shown in the table that follows: TABLE - CABIN INTERPHONE - FLIGHT-ATTENDANT HANDSET/CRADLE REF. BUTTON FUNCTION 1 PA Allows communication between flight attendants and passengers. 2 ATTND Allows communication between the flight attendants. 3 PILOT Allows communication between flight attendants and cockpit crew. 4 EMER PILOT Allows communication between the flight attendants and pilot in emergency mode.
Operation The handset is held on the cradle by means of magnetic latches. To operate the handset, remove it from the cradle and push one of the buttons (PA, ATTND, PILOT and EMER PILOT). The operation of the attendant handset has four different modes, as described: • PA mode: When the PA pushbutton is pushed, the PA LED comes on on each handset cradle. After that, by pressing the PTT button on the handset your voice comes through the cabin loudspeakers. • ATTND mode: When the ATTND pushbutton is pushed, a single chime comes through the cabin loudspeakers and the ATTND LED comes on and flashes on each handset cradle. Communication between the flight attendants is possible by removing any other handset from Issue: June06 Revision: 00
the cradle. • PILOT mode: When the PILOT pushbutton is pushed, a single chime comes through the cockpit loudspeakers. The CAB pushbutton light comes on and flashes on the DAP (Digital Audio Panel)s, the PILOT LED comes on and flashes on each handset cradle, and the PILOT CALL light (red) comes on and flashes on the forward and aft attendant light indicator panels. The communication starts by pushing the CAB pushbutton on one of the DAPs. As a result, the CAB pushbutton light becomes stable on (on the other DAPs, the light goes off), the PILOT LED becomes stable on on each handset cradle, and the PILOT CALL light (red) becomes stable on on the forward and aft attendant light indicator panels. • EMER PILOT mode: When the EMER PILOT pushbutton is pushed, a triple chime comes through the cockpit loudspeakers. The EMER pushbutton light comes on and flashes on the DAPs, the EMER PILOT LED comes on and flashes on each handset cradle, and the pilot emergency call light (red) comes on and flashes on the forward and aft attendant light indicator panels. The communication starts by pushing the EMER pushbutton on one of the DAPs. As a result, the EMER pushbutton light becomes stable on (on the other DAPs the light goes off), the EMER PILOT LED becomes stable on on each handset cradle, and the pilot emergency call light (red) becomes stable on on the forward and aft attendant light indicator panels. To stop the communication put the handset back on its cradle. All the related lights go off.
FOR TRAINING ONLY Reproduction Prohibited
Chapter 44-11
Page 9
Figure 5: PACIC System Schematic 1
B
MONUMENT ( REF.)
1 MONUMENT ( REF.) PA
ATTND
PILOT
EMER PILOT
A P
T T
A
B
A F OR WA R D
F OR WA R D MON UME NT
AV I ONI CS
COMP T
FUS EL A GE
1
R C VR PT T
HA N DS ET / CR A DL E A S S Y A T T #1 (S DS 44 - 13 ) (MPP 44 - 13 - 01)
L H CBP
MI C
B
V
C OC K PI T
DC E S S B US
SE L
CALL
L GHT I
E ME R
LG I HT
A T T L I GHT CABI N I NPH
P AX
C AB I N
3
5
P A CI C
L OUDS PE A KE R S
(S S M 4 4 11- 80 )
(S S M 4 4 12 -8 0 )
P A L I GHT C E NT E R F US I I I MON UME NT R C VR
1
PT T 28
VDC
HA N DS ET / CR A DL E A S S Y A T T #2 (S DS 44 - 13 ) (MPP 44 - 13 - 01)
Issue: June06 Revision: 00
MI C
B
V
FOR TRAINING ONLY Reproduction Prohibited
SE L
Chapter 44-11
Page 10
170/190 MAINTENANCE TRAINING MANUAL
CD Player CD PLAYER. The CD player has the capability of playing one 4.7 in CD at a time, and allows complete control of track selection and play. It features track search, track repeat, advance, reverse, as well as pause functions. If the power fails, a low-current keep-alive signal keeps the memory. Operation The CD player is fully operated by means of buttons on the front panel. To start the operation of the CD player, push the PWR button and insert a CD into the CD slot with the label side up, until the unit automatically pulls it and the display shows the message LOAD. The unit starts to play the first track of the CD. The display shows the track number and its elapsed time. While the CD plays: • Push the VOL UP button to increase the volume. • Push the VOL DOWN button to decrease the volume. • Push the SEL button to select the functions that follow: VOL/BAS/ TRE/BAL/FAD. Use the UP and DOWN buttons to set them. • Push the DWN button to return to the start of the current track. Push it again to go back to the start of the previous track. • Push the UP button to go to the start of the next track. • Push the R/R button to start the repeat mode. The REP indicator shows on the display and the track plays again and again. Push the R/R button to cancel the mode. • Push and hold the R/R button for more than two seconds for random selection of the available tracks. The R indicator shows on the display. • Push the EJECT button to eject the CD.
• Push the PWR button to turn off the unit.
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Chapter 44-11
Page 11
Figure 6: CD Player System Schematic
A
B
AICD III
AVIONICS INNOVATIONS
MUTE
PWR SOURCE
LOUD
DISP
BAND
R/R DWN
UP
1
2
3
4
5
6
TUNE/TRACK
VOL SEL
LOC
VOL
MONO
B
A
RH CBP
C OC K PI T
DC B US
2 F OR WA R D
P R A/ MUS I C
FUS EL A GE
GA L L E Y 2 (S S M 2 5- 3 1- 8 2)
5
F OR WA R D
AV I ONI CS
P AX
COMP T
C AB I N
GA L L E Y C ONT ROL MOD UL E (S DS 25 - 36 ) C D P L AY E R C D P L AY E R
2 8 V DC 2
Issue: June06 Revision: 00
(S DS 44 - 2 1 ) (MPP 44 - 2 1 - 0 1)
A UD I O
P A SS E NGE R A DDR E S S AND C A B I N IT N E RP HO NE C ON T R OL L E R (P AC I C )
A UD I O
L OUDS PE A KE R (S S M 4 4 12 -8 0 )
B
FOR TRAINING ONLY Reproduction Prohibited
(S S M 4 4 11- 80 )
Chapter 44-11
Page 12
170/190 MAINTENANCE TRAINING MANUAL
44-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 2 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 11/16/2004 ¦ 44-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 44 CABIN SYSTEMS ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 3 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 08/26/2005 ¦ 44-2 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 44 CABIN SYSTEMS ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Alternate, normal and ¦ emergency procedures and/or ¦ operating restrictions are ¦ established and used, and ¦ b) Flight attendant alerting ¦ system (audio and visual) ¦ operates normally.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 13-00 Cabin Service ¦ Interphone System
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ NOTE: ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦
C ¦ 1 ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) PA not required by FAR, and ¦ b) Alternate, normal and ¦ emergency procedures and/or ¦ operating restrictions are ¦ established and used.
¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ NOTE: ¦ ¦
¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
B ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
Any station function(s) that operate normally may be used.
Any station function(s) that operate normally may be used.
¦ 12-01 Passenger Cabin ¦ Speakers ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
May be inoperative provided any seat from which a passenger cannot clearly hear a passenger address announcement is blocked and placarded "DO NOT OCCUPY".
¦ ¦ ¦
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ Passenger Address System is ¦ considered inoperative.
¦ ¦ ¦
¦ 12-02 Lavatory Speakers C ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
Any station function(s) that operate normally may be used.
¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Cabin to cabin interphone ¦ functions operate normally ¦ on at least fifty percent of ¦ the cabin handsets, and ¦ b) Alternate communication ¦ procedures between the ¦ affected flight attendant ¦ stations are established and ¦ used.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ NOTE: ¦ ¦
¦ ¦ ¦
(O)May be inoperative provided alternate communication procedures between the affected flight attendant stations are established and used.
Any station function(s) that operate normally may be used.
MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 13-07 Flight Attendant ¦ Alerting System
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: ORIGINAL ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 12/16/2003 ¦ 44-3 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 44 CABIN SYSTEMS ¦ ¦ ¦
¦ 13-01 Flight Attendant ¦ Handsets ¦ ¦ ¦ ¦ ¦ ¦ ¦
B ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) Fifty percent of cabin ¦ handsets operate normally, ¦ and ¦ b) Alternate communication ¦ procedures between the ¦ affected Flight Attendants ¦ station(s) are established ¦ and used.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ NOTE 1: ¦ ¦ ¦ ¦
An operative handset at an inoperative flight attendant seat shall not be counted to satisfy the fifty percent requirement.
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ NOTE 2: ¦ ¦
Any handset function(s) that operate normally may be used.
¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
1) Flight B ¦ 1 Attendant Call ¦ Light System ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
¦ ¦
¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (O)May be inoperative provided: ¦ a) PA system operates normally, ¦ b) If affected light is used ¦ for lavatory smoke detector ¦ alerting, an alternate ¦ lavatory smoke alert (audio ¦ or visual) is installed and ¦ operates normally, and ¦ c) Alternate procedures for ¦ contacting flight attendants ¦ are established and used.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ NOTE 1: ¦ ¦ ¦
Passenger to attendant call system is considered a passenger convenience item.
¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ NOTE 2: ¦ ¦
Any visual alerting system function(s) that operates normally may be used.
¦ (O)May be inoperative provided: ¦ a) PA system operates normally, ¦ b) If affected chime is used ¦ for lavatory smoke detector ¦ alerting, an alternate ¦ lavatory smoke alert (audio ¦ or visual) is installed and ¦ operates normally, and ¦ c) Alternate procedures for ¦ contacting flight attendants ¦ are established and used.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ ¦ ¦ ¦
¦ NOTE 1: ¦ ¦ ¦
Passenger to attendant call system is considered a passenger convenience item.
¦ ¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
¦ NOTE 2: ¦ ¦
Any audio alerting system function(s) that operates normally may be used.
46-00 Information System General General The Embraer 170 is equipped with a Passenger Cabin Information / Boarding Music System that stores passenger briefing messages in different languages and music.The system is installed in the galley 2. Currently, there are two different options for the Passenger Cabin Information/Boarding Music system: • Option1: PBS400 and CD Player; • Option2: PBS600.
Issue: June06 Revision: 00
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Chapter 46-00
Page 1
Figure 1: General
MESSAGE LANGUAGE
VOLUME
SELECT
PBS600 PLAY/PAUSE PWR
Galley 2 Issue: June06 Revision: 00
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170/190 MAINTENANCE TRAINING MANUAL
46-40 Pax- Cabin Information and prerecorded Announcements The PBS400 System Let's have a closer look at the PBS 400 System. It is comprised of two separate units: The Control Head and the Remote Computer, both electrically interconnected and located in Galley 2. The PBS 400 Control Head contains all control buttons and display messages for system operation. Here the required languages and messages can be selected and played over the PA system (Passenger Address). The AICD III CD player can play one compact disc at a time. It also is installed at galley 2.
Issue: June06 Revision: 00
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Chapter 46-40
Page 1
Figure 1: PBS System
PBS - 400
1 PWR
PLAY
BRIEFING
2
LANGUAGE 3
4
Galley 2 Issue: June06 Revision: 00
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Page 2
170/190 MAINTENANCE TRAINING MANUAL
Self test Pressing and releasing the "ON" button turns on the PBS 400. The LED momentarily displays the message 'TESTING' to indicate that the self-test has been initiated. During the power up process, the computer performs a self test. It also flashes several messages in succession at approximately twosecond intervals: Filename, Revision Date, Version Number, Airplane type, Data Test. A successful test is indicated by a message on the LED (Light Emitting Diode).
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Chapter 46-40
Page 3
Figure 2: Self test messages
PBS - 400
1 PWR
PLAY
BRIEFING
2
LANGUAGE 3
4
Messages - TESTING - Filename - Revision date - Version number - Airplane type - Data test - TEST OK
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Page 4
170/190 MAINTENANCE TRAINING MANUAL
Operation The language keys, labled 1,2,3 and 4 are used to select and deselect the required languages. The PBS400 can hold up to 4 languages. Languages that are played are called 'active' languages. Pressing a language key causes that language to be active or inactive. A character appearing in the right side of the LED display indicates active languages. (For example, pressing the '1' button once causes the 'E' (English) symbol to appear, then pressing it again causes the 'E' to disappear.) The left-most active language symbol on the display indicates the language that will be played first, and the right-most symbol the language that will be played last. To exchange the order of the languages, pressing '1' to deselect English, then pressing '1' again to reselect English in different order.
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Chapter 46-40
Page 5
Figure 3: operation
PBS - 400
1
PWR
land
ef
PLAY
English
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LANGUAGE 3
BRIEFING
French
FOR TRAINING ONLY Reproduction Prohibited
2
4
Language selection key
Chapter 46-40
Page 6
170/190 MAINTENANCE TRAINING MANUAL
Operation (continued) The UP/DOWN arrow is used to select passenger briefings, by scrolling through the list of available message titles. Note that the message titles may be abbreviated. For example: TAKEOFF is abbreviated as T/O. To start the briefing press PLAY. The active language indications will be replaced by the message PLAY. The voice message will then play. The PBS400 will play the message in the first active language, than automatically start the next active language.
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Chapter 46-40
Page 7
Figure 4: Operation (continued)
Announcement select key
PBS - 400
1 PWR
T t/o
ef
PLAY
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LANGUAGE 3
BRIEFING
FOR TRAINING ONLY Reproduction Prohibited
2
4
Chapter 46-40
Page 8
170/190 MAINTENANCE TRAINING MANUAL
Operation (continued) To interrupt the message in process, press the PLAY button again. The PLAY indication will extinguish. The PBS will return to the beginning of the sentence that it was speaking when interrupted, and wait. Pressing the PLAY button again will restart the message from the beginning of the sentence. To restart the briefing from the beginning, use the UP/DOWN arrow to select a different message, than reselect the desired message. This process 'reinitializes' that message. Then press the play button. The PBS 400 will be automatically interrupted if a member of the flight deck crew or cabin crew presses a push-to-talk button on the any of the interphones for PA (passenger adress) announcements. When the push-to-talk button on the interphone system is released, the PBS400 will resume play from the beginning of the sentence that was interrupted.
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Chapter 46-40
Page 9
Figure 5: Operation (continued)
Play / Pause Announcement key
PBS - 400
1 PWR
land
play
PLAY
LANGUAGE 3
BRIEFING
2
4
keys to reselect complete announcement after interruption Issue: June06 Revision: 00
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Page 10
170/190 MAINTENANCE TRAINING MANUAL
Operation (continued) Once a manual message has been completed, the computer will not allow the message to be given again simply by pressing the PLAY button. This feature prevents accidental sending of a message once it has been given. To repeat a message, select a different message using the UP/DOWN keys, than reselect the desired message. This process initializes the message. Press PLAY to start the briefing.
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Chapter 46-40
Page 11
Figure 6: Operation (continued)
After interruption by manual announcement.....
PBS - 400
1 PWR
land
eG
PLAY
LANGUAGE 3
BRIEFING
2
4
up / down keys
PBS - 400
1 PWR
land
play
BRIEFING
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PLAY
2
LANGUAGE 3
4
Chapter 46-40
Page 12
170/190 MAINTENANCE TRAINING MANUAL
Operation of the CD- player The CD Player can be operated via front panel buttons. Press the "PWR" button to turn on the CD payer. Insert the CD into the CD slot. "LOAD" will be displayed, and the CD player begins to play from the first track on the disk. Press the "DWN" button to return to the start of the current track. Press it again to go back to the start of the previous track. Press the "UP" button to advance to the beginning of the next track. Press the "R/R " button to repeat the track currently played. Use the eject button to eject and remove the CD from the slot.
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Chapter 46-40
Page 13
Figure 7: Operation of the CD- player
eject key CD in AICD III
AV I O N I C S I N N O VAT I O N S
PWR
UP
MUTE
load DWN
1
2
3
4
5
TUNE / TRACK
DISP
VOL
R/R
SEL
6
VOL LOCAL
track selection keys
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Page 14
170/190 MAINTENANCE TRAINING MANUAL
The PBS600 Now let's look at the PBS 600. The audio entertainment system is installed on the upper outboard area of the forward galley, and provides the flight crew and passengers with safety and flight information. The Prerecorded Announcement Assembly Unit, is a computer-controlled digital voice multi-language single-channel passenger briefing system. It provides a convenient and consistent method of informing passengers of important messages such as safety instructions, takeoff, landing, over water, and other airline information, and also provides the passengers and flight crew with music. In addition, the unit can deliver up to seven hours of pre-recorded music. A real human voice is digitally recorded and stored directly into the system internal memory. The messages can be spoken in up to sixty-three different languages. As an option, the unit can be configured to play pre-recorded safety and informational messages in conjunction with pilot or copilot actions, such as illuminating a FASTEN SEAT BELT or NO SMOKING cabin warning light.
Issue: June06 Revision: 00
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Chapter 46-40
Page 15
Figure 8: The PBS600
MESSAGE
MESSAGE Pretakeoff Takeoff Seatbelt On Seatbelt Off Terminating Fli Intermediate Fli
LANGUAGE
VOLUME
SELECT
ENGLISH FRENCH GERMAN ITALIAN SB / OFF CUED
PBS600 PLAY/PAUSE PWR
Safety announcements Music COMPUTER CONTROLLED DIGITAL VOICE MULTI LANGUAGE SINGLE CHANNEL PASSENGER BRIEFING
Optional
SYSTEM
Issue: June06 Revision: 00
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Chapter 46-40
Page 16
170/190 MAINTENANCE TRAINING MANUAL
Operation PBS600 To turn the prerecorded announcement assembly ON, the power button has to be pressed and released. A self-test will be performed, and a selection menu will appear on the Liquid Crystal Display. By pushing and releasing the LANGUAGE button, a list of up to 63 different languages is displayed. To scroll through the list, the UP and DOWN selection arrow buttons have to be used. To select one or more languages, the SELECT button has to be pushed. However, only four of these languages may be selected at a time. To de-select a language, use the UP and DOWN selection arrow buttons and then push the SELECT button. The MESSAGE button has to be pushed and released to display a list of message groups. It is possible to scroll through the list by pushing the UP and DOWN arrow buttons. For selecting a message, the SELECT button has to be used. The PLAY/PAUSE button is used to start or to stop a message. The message will be re-started from its beginning if it was stopped and re-started. The message is played in each of the languages in the order in which they were selected. The volume will be adjusted by pushing the VOLUME UP and DOWN arrow buttons while the message is playing. To skip quickly forward or backward through the available messages in a message group, use the BACK and NEXT arrow buttons. To turn the unit off, push and release the power button.
Issue: June06 Revision: 00
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Chapter 46-40
Page 17
Figure 9: Operation PBS600
MESSAGE
MESSAGE Pretakeoff Takeoff Seatbelt On Seatbelt Off Terminating Fli Intermediate Fli
LANGUAGE
Message / Language select keys
VOLUME
SELECT
ENGLISH FRENCH GERMAN ITALIAN SB / OFF CUED
PBS600 PLAY/PAUSE PWR
Previous / Next select keys
Issue: June06 Revision: 00
FOR TRAINING ONLY Reproduction Prohibited
Power On Self test
Chapter 46-40
Page 18
170/190 MAINTENANCE TRAINING MANUAL
46-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: ORIGINAL ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 12/16/2003 ¦ 46-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 46 INFORMATION SYSTEMS ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 41-00 Prerecorded ¦ *** Passenger ¦ Announcement ¦ System
C ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used. ¦
¦ ¦ ¦ ¦
¦ ¦
D ¦ ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
33-10 Lighting The Flight compartment lighting The flight compartment lighting system provides lighting to the work area, panels and instruments, and consists of the following sub-systems: • The cockpit lights system, which provides beam ambient lighting; used on the side walls, seats, and floor of the crew station and observer area. • The instrument and control panel lights system, which provides lighting for instruments, panels, and push buttons. • The flood/storm lights system, which provides a proper lighting level in the cockpit for the instruments and assures instrument readability.
last update: Dec06
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Chapter 33-10
Page 1
Figure 1: Flight compartment lighting
last update: Dec06
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Chapter 33-10
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170/190 MAINTENANCE TRAINING MANUAL
The cockpit lights The cockpit lights include the following components: • Two dome lights located on the cockpit ceiling, which are controlled by a switch on the overhead panel. Power to the dome lights is supplied by the Essential DC Bus No 3. This permits cockpit lighting in electrical emergency conditions. • The pilot, copilot and observer reading lights, installed on the ceiling. They help the flight crew to read maps, checklists and manuals. The lighting intensity level can be adjusted by clockwise or counterclockwise movements of the reading lights inner bezel. The units permit a light beam orientation, zoom, and full movement in any direction by turning the outer bezel. • The pilot's and co-pilot’s chart holder assembly each have an associated chart light for adequate illumination. They are located over the left and right side windows. The light beams can be deflected up to 20 degrees from vertical axis, which provides a total movement of approximately 40 degrees. Two on/off variable-intensity potentiometers located on the pilot's and co-pilot’s glare shield panels control the chart holder lights brightness. Turning the knob to the full counter-clockwise position turns the light off. Reading plane area is adjustable by rotating the chart lights bezel.
last update: Dec06
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Chapter 33-10
Page 3
Figure 2: The cockpit lights
last update: Dec06
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Chapter 33-10
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170/190 MAINTENANCE TRAINING MANUAL
Operation The DC ESS BUS 3 supplies 28 VDC (Volt Direct Current) to the dome lights, which can be controlled by the DOME LIGHTS switch.The dome lights can also be controlled by an automatic courtesy light logic provided that the aircraft is in the ground service configuration and the DOME LIGHTS switch is set in the ON position. The DC BUS 2 supplies 28 VDC to the reading lights, which are controlled by rotation of the bezel. The DC BUS 1 supplies 28 VDC to the chart lights, which are controlled by the CHART knob located in the LH and RH lighting control panels.Turning the CHART knob fully clockwise causes the lights to have a normal brightness.The lights have minimum brightness if the knob is turned fully counterclockwise. The chart lights provide directional control of the light beam and can be used to supplement the reading lights if desired. In the DIM position, the lights come on with minimum brightness.In the BRT position, the light comes on with normal brightness.
last update: Dec06
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Chapter 33-10
Page 5
Figure 3: Cockpit lights
E
D
C
EICA
C FLO CHAR
B
B
S
MFD PFD
BRT PTT
OD/
STO
RM
OFF
BRT
DIM
BRT
BRT
DIM
COCKPIT LIGHTS
OFF
F
A
BRT
T
DIM
A
MAIN PNL
C
B
OFF
OVHD PNL
BRT
OFF
PEDESTAL
BRT
OFF
BRT
DOME LIGHT ON
ANNUNCIATORS TEST
OFF
COCKPIT CEILING PNL
COCKPIT CEILING PNL PILOT DOME LIGHT
(SDS 33-11 (MPP 33-11-01)
COCKPIT CEILING PNL
COPILOT READING LIGHT
COPILOT DOME LIGHT
E
(SDS 33-11 (MPP 33-11-01)
E
D C
(SDS 33-11 (MPP 33-11-01)
COCKPIT CEILING PNL
COCKPIT CEILING PNL PILOT READING LIGHT
OBSERVER READING LIGHT
C
C
(SDS 33-11) (MPP 33-11-01) LH WINDOW
GLARESHIELD
OVERHEAD PANEL
(SDS 33-15) (MPP 31-13-01)
PILOT CHART LIGHT
RH WINDOW
GLARESHIELD (SDS 33-15) (MPP 31-13-01)
COCKPIT LIGHTS PANEL
(SDS 33-11) (MPP 33-11-01)
COPILOT CHART LIGHT
DOME LIGHT SW (SDS 33-10) (MPP 33-10-01)
B
(SDS 33-11) (MPP 33-11-05)
F
B D
A
(SDS 33-11) (MPP 33-11-05)
COCKPIT
LH CB PANEL
DC BUS 1
COCKPIT CHART LTS
5
SSM 33-26-80 DC ESS BUS 3
COCKPIT DOME LTS
RH CB PANEL
COCKPIT
5 COCKPIT READING LTS
DOME LIGHT CTRL RLY
last update: Dec06
F
FOR TRAINING ONLY - Reproduction Prohibited
E
5
DC BUS 2
Chapter 33-10
Page 6
170/190 MAINTENANCE TRAINING MANUAL
Cockpit lights overhead panel Integrated in the overhead panel is the lights control panel.It associates with the on/off variable-intensity potentiometers.Turning these potentiometers in the clockwise direction brightens the background lights of the main panel, the overhead panel and the pedestal.As a discrete function of the overhead panel potentiometer, the annunciator lights can be set to daylight (bright) and night-time (dim) illumination.Turning the overhead panel lights to the off position will automatically set the annunciator lights to bright intensity. All three potentiometers are connected to a multi channel dimmer. The multi channel dimmer varies its output voltage from 0 to 5.5 volts depending on the dimmer position. Output power is then sent to the respective lights. The annunciator lights can be tested by pushing the annunciator lights test push switch.
last update: Dec06
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Chapter 33-10
Page 7
Figure 4: Cockpit Lights overhead panel
Multi Function Dimmer
The multichannel dimmer varies its output voltage from 0 to 5.5 volts depending on the dimmer position.
last update: Dec06
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Chapter 33-10
Page 8
170/190 MAINTENANCE TRAINING MANUAL
Components DIMMERS There is one electronic dimmer located in the forward avionics compartment that provides the necessary dimming function for the illumination of the pushbuttons, overhead panel, pedestal, and main panel. Each channel has an independent 28 VDC power source and circuitry designed to provide linear control of LED illumination levels. The output of each light dimmer channel may be controlled by means of a single-turn potentiometer.
last update: Dec06
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Chapter 33-10
Page 9
Figure 5: Cockpit lights
B A
B
A
COCKPIT LIGHTS OVERHEAD PANEL COCKPIT
LH CB PANEL
MAIN PNL
OVHD PNL
PEDESTAL
COCKPIT LIGHTS PANEL MAIN PNL SW OFF
DC BUS 1 MAIN PANEL
BRT
OFF
BRT
7.5 OFF
BRT
C
(SDS 33-10) (MPP 33-10-01)
OFF
BRT
DOME LIGHTS ON
ANNUNCIATORS TEST
OFF
C
FWD FUSELAGE
DIMMER SSM 34-11-80
(SDS 33-12) (MPP 33-12-01)
CH 1
B
MAIN PANEL LH LIGHT CONTROL PANEL (SSM 33-15-80))
last update: Dec06
MAIN PANEL AUTOBRAKE PNL LTS (SSM 32-41-80)
MAIN PANEL REVERSIONARY PNL LTS (SSM 31-61-80)
MAIN PANEL G/S INHIBIT PNL LTS (SSM 34-41-80)
FOR TRAINING ONLY - Reproduction Prohibited
MAIN PANEL
MAIN PANEL
MAIN PANEL
REVERSIONARY PNL LTS
TERRAIN INHIBIT PNL LTS
RH LIGHT CONTROL PANEL
(SSM 31-61-80)
(SSM 32-44-80)
(SSM 33-15-80)
Chapter 33-10
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170/190 MAINTENANCE TRAINING MANUAL
Glare shield panel illumination The flood or storm lights provide sufficient lighting intensity to illuminate the main cockpit panel during lighting conditions, and to avoid blinding of the pilots and co-pilots eyes. The flood lights are located under the main panel and controlled by an on/off variable intensity potentiometer on the glare shield panel. The pilot's PFD, MFD and EICAS display brightness controls are located on the left glare shield panel, while the co-pilot's PFD, MFD and standby instruments brightness is controlled from the right glare shield panel.
last update: Dec06
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Chapter 33-10
Page 11
Figure 6: Glare shield panel illumination
last update: Dec06
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Chapter 33-10
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170/190 MAINTENANCE TRAINING MANUAL
33-20 Passenger Compartment Lights Introduction The passenger compartment lighting system provides illumination for the cabin. It consists of: • • • • • •
the passenger cabin lights system, the passenger warning signs system, the reading and attendant call lights system, the courtesy/stair lights system, the lavatory lights system and the galley lights system.
last update: Dec06
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Chapter 33-20
Page 1
Figure 1: Passenger compartment lights
last update: Dec06
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Chapter 33-20
Page 2
170/190 MAINTENANCE TRAINING MANUAL
The passenger cabin lights The passenger cabin lights provide general illumination of the cabin and include the ceiling lights, mounted above the passenger service unit structure and the side wall lights mounted along the side walls. Control is provided by 4 switches located on the forward and aft flight attendants panel. Two of the switches are used to turn the lights on and off, and the other two switches are used to control the brightness of the lights. Ceiling Ballasts provide electrical power and regulate the ceiling and side wall lights. These ballasts are mounted throughout the cabin behind the passenger service units.
last update: Dec06
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Chapter 33-20
Page 3
Figure 2: The passenger cabin lights
last update: Dec06
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Chapter 33-20
Page 4
170/190 MAINTENANCE TRAINING MANUAL
Ceiling/Sidewall light ballasts The ceiling and sidewall light electronic ballasts control the power supply to the fluorescent ceiling lights.Besides reducing flickering, these ballasts are provided with a power factor controller that reduces harmonic distortion.
last update: Dec06
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Chapter 33-20
Page 5
Figure 3: Cabin lights
1
E
C
MONUMENT ( REF.)
C
CABIN LIGHTING
ON
SIDEWALL
A
BRIGHT DIM
B
A 2ND RH BAG BIN
LAMP / BALLAST MID AVIONICS COMPT
(SDS 33-21) (MPP 33-21-11)
LICC
C
B
1ST RH BAG BIN
3RD RH BAG BIN
LAMP / BALLAST
E
LAMP / BALLAST
E
(SDS 33-21) (MPP 33-21-11)
4TH RH BAG BIN LAMP / BALLAST
E
(SDS 33-21) (MPP 33-21-11)
(SDS 33-21) (MPP 33-21-11)
F
5TH RH BAG BIN
6TH RH BAG BIN
LAMP / BALLAST
E
(SDS 33-21) (MPP 33-21-11)
7TH RH BAG BIN
LAMP / BALLAST
E
(SDS 33-21) (MPP 33-21-11)
8TH RH BAG BIN
LAMP / BALLAST
E
(SDS 33-21) (MPP 33-21-11)
9TH RH BAG BIN
LAMP / BALLAST
E
(SDS 33-21) (MPP 33-21-11)
LAMP / BALLAST
E
E
(SDS 33-21) (MPP 33-21-11)
SIDEWALL LTS
AC GND SVC BUS
PHASE C
PHASE B
MID AVIONICS COMPT
FWD AVIONICS COMPT
PHASE A 5
REAR PLATE
REAR PLATE
MID AVIONICS COMPT
DISCRETE IN / OUT MODULE (SLOT 11)
MONUMENT (REF.)
(SSM 24-61-80)
SPDA 2
(SSM 24-61-80)
SPDA 1
REAR PLATE
ON IND
BRT IND
DIM IND
DIM/BRT SW
ON IND
DIM IND ON/OFF SW
DIM/BRT SW
C
(SDS 33-21) (SWPM 20-32-01)
REAR PLATE
A MONUMENT CAN BE A LAVATORY OR A CLOSET
last update: Dec06
AFT FLIGHT ATT PANEL (SDS 33-21) (MPP 25-25-02)
DISCRETE IN/OUT MODULE (SLOT 11)
LAMP / BALLAST
1
BRT IND
FWD FLIGHT ATT PANEL (SDS 33-21) (MPP 25-25-01)
C
DISCRETE I/O MODULE (SLOT 9)
AFT LAVATORY
(SSM 24-42-50)
DISCRETE IN/OUT MODULE (SLOT 11)
DISCRETE IN / OUT MODULE (SLOT 11)
DC POWER MODULE (SLOT 4)
ON/OFF SW
1
SPDA 2 (SSM 24-61-80)
SPDA 1 (SSM 24-61-80)
SIDEWALL LIGHTS RLY
(SDS 33-21) (MPP 33-21-15)
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 2ND LH BAG BIN
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 3RD LH BAG BIN
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 4TH LH BAG BIN
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 5TH LH BAG BIN
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 6TH LH BAG BIN
FOR TRAINING ONLY - Reproduction Prohibited
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 7TH LH BAG BIN
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15) 8TH LH BAG BIN
LAMP / BALLAST
F
(SDS 33-21) (MPP 33-21-15)
F
9TH LH BAG BIN
Chapter 33-20
Page 6
170/190 MAINTENANCE TRAINING MANUAL
The passenger warning signs The passenger warning signs provide the passengers and flight attendants with the following signs: • • • •
NO SMOKING, FASTEN SEATBELTS, RETURN TO SEAT, and LAVATORY OCCUPIED.
The system comprises: • The no smoking switch located on the overhead panel, which controls the no smoking signs located in the passenger service units; • the fasten seat belt switch, also located on the overhead panel, which controls the fasten seat belt signs located in the passenger service units, and the return to seat sign in the lavatory; and the lavatory door latch engaged switch, which controls the lavatory occupied signs. Note that the NO SMOKING and FASTEN SEATBELT signs will illuminate automatically in case of depressurization or the cabin altitude exceeding 14,000 ft.
last update: Dec06
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Chapter 33-20
Page 7
Figure 4: The passenger warning signs
The fasten seatbelt switch controls the fasten seatbelt signs located in the passenger service units, and the return to seat sign in the lavatory.
Note that the NO SMOKING and FASTEN SEATBELT signs will illuminate automatically in case of depressurization or the cabin altitude exceeding 14,000 f
last update: Dec06
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Chapter 33-20
Page 8
170/190 MAINTENANCE TRAINING MANUAL
The reading and attendant panel The reading and attendant call lights system provides adequate lighting for reading, and information for attendants regarding passenger or pilot needs. The attendant call panel contains the attendant call lights, which provide visual indication to the attendant when a call from the pilots or passengers is received. Four blue zonal lights, distributed along the cabin ceiling, help the Flight Attendant locate the call. The ATDT call lights provide a visual indication to ATDT when there is a call from the flight crew or passengers. Each type of call to the flight ATDT is identified by a specific colour. For all types of call there is just one colour, that could be orange, blue, amber, red and green. • Orange: when a pax presses the ATDT call SW located inside the lavatory. • Blue: when a pax in the cabin area presses the ATDT call SW on the PSU located above the seat. • Amber: when the pilot does not want to be disturbed. This light is controlled by a SW installed on the overhead panel in the cockpit, and is designated to illuminate the sterile light. • Red: when the pilot makes an emergency call to the flight ATDT from the cockpit. • Green: when the pilot calls the flight ATDT from the cockpit. Test switches installed on the attendant panel are used to check the PSU reading lights, the attendant call lights and the attendant call annunciator lights. Attendant call switches are located on each passenger service unit and in the lavatory. Attendant call annunciator lights are also located in the passenger service units. The PSU switch assemblies on the passenger service units are used to control PSU reading lights that provide adequate lighting for all passenger seats. There are also reading lights mounted over each flight attendant seat. They are individually controllable by switches.
last update: Dec06
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Chapter 33-20
Page 9
Figure 5: The reading and attendant panel
ER R ST
PSU
PSU
last update: Dec06
LAV
MER
E ATTDT
ME TDT E T A LAV
STER
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Chapter 33-20
Page 10
170/190 MAINTENANCE TRAINING MANUAL
The courtesy lights The courtesy lights provide lighting for: • • • • •
forward and rear main and service door entrances, the air stairs, the cockpit step between the cockpit and the cabin, the cockpit dome lights and the DC ceiling lights mounted on the overhead bins.
Courtesy lights are controllable by a switch mounted on the flight attendant panels. The switch gives the crew the option of having the courtesy lights in Off or Auto mode. Under normal operating conditions, the DC lights are controlled by the cabin lighting system. When normal aircraft power is not available, for example on ground with the APU not running, it is still possible to use the courtesy lights and DC ceiling lights. In this case, the Hot Bus provides power to these lights. To conserve battery power, these lights operate on a five minute cycle. A reset button located on the flight attendant panels, and as an option, on the cockpit pedestal panel, allows an additional 5 minutes of lighting every time the switches are pressed.
last update: Dec06
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Chapter 33-20
Page 11
Figure 6: The courtesy lights system
The courtesy lights provide lighting for: Forward and rear main and service door entrances The airstairs The cockpit step between the cockpit and the cabin The cockpit dome lights The DC ceiling lights mounted on the overhead bins
Cockpit step
FWD service door
DC ceiling lights
Rear service door
Cockpit dome light
FWD main door
Rear main door
Air stairs
AIr stairs
Courtesy lights
last update: Dec06
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Chapter 33-20
Page 12
170/190 MAINTENANCE TRAINING MANUAL
Operation NORMAL OPERATION Under normal operating conditions, the courtesy-lights are controlled by the ground service.Courtesy-lights are controlled by a door-mounted switch and a rotary switch mounted on the flight attendant panels.The door-mounted microswitch is responsible to control the related courtesy lights and DC (Direct Current) lights (forward and aft aircraft part).The rotary switch gives the aircraft crew the option of having the courtesy lights in the OFF mode or AUTO mode. HOT BATT BUS OPERATING Under the situations where ground service power is not available it is still possible to have operation of the courtesy lights and DC ceiling lights.During these conditions, the HOT BATT BUS provides power to these lights.A fiveminute time allowance will be allowed to operate these lights for battery power conservation.Reset buttons located on the flight attendant panels will allow an additional 5 minutes of lighting every time the switches are pressed.In the HOT BATT BUS power condition, the DC cabin light will function in DIM mode to conserve aircraft battery power.
last update: Dec06
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Chapter 33-20
Page 13
Figure 7: Courtesy lights
A
F
E
E B
B
F E
A
CENTRAL FUSELAGE III MAIN ENTR COURTESY LT
MIDDLE AVIONICS COMPT
LICC
E
COCKPIT
LH CKT BRKR PNL
FWD AVIONICS COMPT DC GND SVC BUS
HOT BAT BUS 1
(SDS 33-26) (MPP 33-26-01)
LH RELAY SUPPORT
AFT L MAIN CLG PNL
AFT COURTESY LTS TIMER RELAY COURTESY LIGHTS
7.5
COURTESY
F
7.5
CENTRAL FUSELAGE III
SHEET 02 SVC DOOR COURTESY LT SHEET 02
C
E (SDS 33-26) (MPP 33-26-05) AFT R MAIN CLG PNL FWD FUSELAGE
(SDS 33-26) (MPP 33-26-05) FWD GALLEY AREA CLG PNL
A
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Chapter 33-20
Page 14
170/190 MAINTENANCE TRAINING MANUAL
The lavatory lights system The lavatory lights system provides ceiling and side wall lighting in the lavatory area. The system includes a dome light installed in the ceiling panel, a ballast located on top of the lavatory ceiling panel, and two light assemblies. When electrical power is applied to the aircraft, the ceiling lights come on in dim mode; as long as the toilet door is not latched the lights remain dimmed. If the toilet door is closed and latched, the ceiling lights, the cabinet and sink lights will come on in the bright mode. At the same time, the lavatory occupied sign will illuminate in the cabin.
last update: Dec06
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Chapter 33-20
Page 15
Figure 8: The lavatory lights system
Ceiling and sidewall lighting Dome light
Ballast
last update: Dec06
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Chapter 33-20
Page 16
170/190 MAINTENANCE TRAINING MANUAL
The galley lights systems The forward and aft galley lights systems provide a direct light source in the galley area for the flight attendants. Each consists of a two lamp housing with fluorescent lamps that receive electrical current from two ballasts which regulate the power output to the lamps. The galley lights are controlled by switches installed on the respective flight attendant panels.
last update: Dec06
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Chapter 33-20
Page 17
Figure 9: The galley lights systems
2 ballasts 2 lamp housing with flourescent lamps
FWD galley last update: Dec06
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Chapter 33-20
Page 18
170/190 MAINTENANCE TRAINING MANUAL
33-30 Baggage and Service Compartments Lights General Lights are installed in service and cargo compartments to provide technical and ground handling personnel with adequate lighting. All of these compartment lights are activated by a micro switch when the respective compartment door is opened.
last update: Jun06
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Chapter 33-30
Page 1
Figure 1: Service and Cargo compartment lights
Micro switch
last update: Jun06
FOR TRAINING ONLY - Reproduction Prohibited
Micro switch
Chapter 33-30
Page 2
170/190 MAINTENANCE TRAINING MANUAL
Cargo bay lights The forward cargo bay has 4 cargo lights and 1 loading light. The aft cargo bay has 3 cargo lights and 1 loading light.There is a manual switch located by each cargo door that can be selected to “AUTO” or "OFF" position. In “AUTO” mode the cargo lights illuminate when the cargo door is opened, and go out when the door closes. The “OFF” mode will turn the lights off regardless of the door position.
last update: Jun06
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Chapter 33-30
Page 3
Figure 2: Cargo bay lights
MICRO SWITCH
MICRO SWITCH
CARGO LIGHT
CARGO LIGHT
LOADING LIGHT LOADING LIGHT
FORWARD CARGO COMPARTMENT last update: Jun06
AFT CARGO COMPARTMENT
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Chapter 33-30
Page 4
170/190 MAINTENANCE TRAINING MANUAL
Operation The operation of the cargo compartment lights is controlled by the position of the door and by the manual switch.To turn the lights on, the manual switch must be in the AUTO mode and the cargo compartment door must be open.To turn the lights off, the door must be closed or the manual switch must be in the OFF mode, in this case, the door position does not interfere.
last update: Jun06
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Chapter 33-30
Page 5
Figure 3: Cargo Compartment lights
C
B
D
E
A
C F
A MID AVIONICS COMPT MID AVIONICS COMPT
REAR PLATE
SPDA 2 (SSM 24-61-80)
DC POWER MODULE (SLOT 7) DC GND SVC BUS
F
DC GND SVC BUS
FWD CARGO LIGHTS
B 15
D
LICC
FWD CARGO COMPT
FWD CARGO COMPT
CARGO LIGHT 3
F
MID AVIONICS COMPT
E FWD CARGO COMPT
CARGO LIGHT 4
(SDS 33-31) (MPP 33-31-01)
F
(SDS 33-31) (MPP 33-31-01)
FWD CARGO COMPT
CARGO LIGHT 2
F
(SDS 33-31) (MPP 33-31-01)
CARGO LIGHT 1
F
(SDS 33-31) (MPP 33-31-01)
RELAY SUPPORT
(SDS 33-30) (SWPM 20-32-01)
last update: Jun06
CARGO BAY LT MANUAL SWITCH
C
C
E
(SDS 33-31) (MPP 33-31-05)
OFF
CARGO DOOR LIGHT SWITCH
(SDS 33-31) (MPP 33-31-09)
FWD CARGO DOOR
FWD CARGO DOOR
AUTO
CLOSE
OPEN
FWD CARGO DOOR
(SDS 33-31) (MPP 33-31-13)
FOR TRAINING ONLY - Reproduction Prohibited
LOADING LIGHT 2
Chapter 33-30
Page 6
170/190 MAINTENANCE TRAINING MANUAL
33-40 Exterior Lights Introduction The exterior lighting system uses high-intensity lights. These lights are used for taxiing, takeoff and landing procedures. They are also used for in-flight orientation and identification of aircraft position, and promotion of the aircraft operator logo. The exterior lights include the following subsystems: • • • • • •
The landing lights, the taxi lights, the navigation lights, the inspection lights, the logo lights and the anti-collision lights.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 1
Figure 1: The exterior lights
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 2
170/190 MAINTENANCE TRAINING MANUAL
The landing lights Two 600 watt landing lights are fitted in the wing leading edges close to the fuselage, and one 600 watt landing light is installed on the nose landing gear assembly. The landing lights receive 28 VAC from transformers installed near the wing roots and nose landing gear bay. They are each controlled by separate switches located on the overhead panel. The landing lights are supplied by the Secondary power distribution assembly or by SPDA 1 and 2. In case of a single failure, two lights can provide enough lighting for a safe landing at night. The landing light installed on the nose landing gear will automatically extinguish when the landing gear is retracted.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 3
Figure 2: The landing lights
Transformers
The landing lights receive 28 VAC from transformers installed near the wing roots and nose landing gear bay.
Two 600W landing lights
External light panel One 600W landing light
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 4
170/190 MAINTENANCE TRAINING MANUAL
Operation LANDING LIGHTS - OPERATION The landing lights are controlled by independent switches, namely LEFT, RIGHT and NOSE, in the cockpit overhead external lights panel. The wing-root landing light transformers interface with SPDA (Secondary Power Distribution Assembly) 2 located in the middle avionics compartment to receive 115 VAC/400 Hz power.The NLG landing light transformers interface with SPDA 1 located in the forward avionics compartment.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 5
Figure 3: Landing lights
E G
D G
C A
EXTERNAL LIGHTS NAV
STROBE
RED BCN
ON
E A
ON
OFF
OFF
LOGO NOSE
G
B
TAXI ON OFF
INSP
SIDE ON
ON
OFF
OFF
LANDING LEFT
F
G
NOSE
RIGHT
ON
ON
OFF
OFF
H
C REAR PLATE
D
REAR PLATE
OVERHEAD PANEL NOSE LDG LT SW
SPDA 1 (SSM 24-61-80)
MID AVIONICS COMPT
LH LDG LT SW
RH LDG LT SW SPDA 2 (SSM 24-61-80)
A FWD AVIONICS COMPT
EXTERNAL LIGHT PANEL AC POWER MODULE AC GND SVC BUS
NOSE LDN GEAR COMPT
DISCRETE IN/OUT MODULE
H
ON
OFF
(SDS 33-40) (MPP 33-40-01)
ON
OFF
ON
OFF
AC POWER MODULE AC BUS 1
DISCRETE IN/OUT MODULE
115 VAC
AC POWER MODULE AC BUS 2
115 VAC
115 VAC
FAIRING
FAIRING
STEPDOWN TRANSFORMER
STEPDOWN TRANSFORMER
G
(SDS 33-41) (MPP 33-41-07)
23 VAC
WING FAIRING FILLET
23 VAC
(SDS 33-41) (MPP 33-41-01)
F
NOSE LDG LIGHT
last update: Dec06
G
(SDS 33-41) (MPP 33-41-05)
FWD FUSELAGE (SDS 33-41) (MPP 33-41-03)
STEPDOWN TRANSFORMER
LH LDG LIGHT
FOR TRAINING ONLY - Reproduction Prohibited
WING FAIRING FILLET
B
G
(SDS 33-41) (MPP 33-41-05)
23 VAC
(SDS 33-41) (MPP 33-41-01)
B
RH LDG LIGHT
Chapter 33-40
Page 6
170/190 MAINTENANCE TRAINING MANUAL
The taxi lights The ERJ 170 has a total of three 450 watt taxi lights. Their location is similar to the landing lights: Two in the wing roots and one on the nose landing gear assembly. 115 volts AC from SPDA 1 and 2 is transformed into 28 volts AC, and then provided to the 3 taxi lights. The taxi lights have two separate switches for appropriate ambient lighting. The two taxi lights in the wing to fuselage fairing have a common switch while the nose wheel taxi light has a separate switch. The nose gear taxi light will automatically extinguish if the nose gear is not in the down locked position.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 7
Figure 4: Taxi lights
SPDA 1, 2 115 volts AC from SPDA 1 and 2 is transformed into 28 volts AC, and then provided to the 3 taxi lights.
Two 450W taxi lights
External light panel One 450W taxi light
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 8
170/190 MAINTENANCE TRAINING MANUAL
Operation TAXI LIGHTS - OPERATION The NLG taxi light is controlled by means of the NOSE switch installed in the cockpit overhead external lights panel. Two wing-root taxi lights are controlled by means of the SIDE switches, one for the right and left sides, installed in the cockpit overhead external lights panel. The NLG taxi lights are operated when the NLG is down and locked only. The wing-root taxi Lights interface with SPDA (Secondary Power Distribution Assembly) 2 located in the middle avionics compartment to receive 115 VAC/400 Hz power. The NLG Taxi Lights interfaces with SPDA 1 located in the forward avionics compartment.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 9
Figure 5: The taxi lights
E G
D G
C A
EXTERNAL LIGHTS
E
NAV
STROBE
RED BCN
ON
A
G
B
ON
OFF
OFF TAXI
LOGO NOSE ON OFF
F
G
ON
OFF
OFF
LANDING LEFT
A
INSP SIDE
ON
NOSE
RIGHT
ON
ON
OFF
OFF
H
C
D
FWD AVIONICS COMPT
MID AVIONICS COMPT
REAR PLATE
REAR PLATE OVERHEAD PANEL
SPDA 1 (SSM 24-61-80)
NOSE TAXI LIGHT SW
TAXI LIGHT SW
AC POWER MODULE AC GND SVC BUS
NOSE LDN GEAR COMPT
DISCRETE I/O MODULE
ON
(SDS 33-40) (MPP 33-40-01)
OFF
ON
AC POWER MODULE AC BUS 1
DISCRETE I/O MODULE
H
115 VAC
AC POWER MODULE AC BUS 2
115 VAC
115 VAC
FAIRING
STEPDOWN TRANSFORMER
FAIRING
STEPDOWN TRANSFORMER
G
(SDS 33-42) (MPP 33-42-07)
FWD FUSELAGE
WING FAIRING FILLET
(SDS 33-42) (MPP 33-42-03)
(SDS 33-42) (MPP 33-42-01)
NOSE TAXI LIGHT
last update: Dec06
STEPDOWN TRANSFORMER
G
(SDS 33-42) (MPP 33-42-05)
23 VAC
F
23 VAC
LH TAXI LIGHT
FOR TRAINING ONLY - Reproduction Prohibited
SPDA 2 (SSM 24-61-80)
EXTERNAL LIGHT PANEL OFF
G
(SDS 33-42) (MPP 33-42-05)
WING FAIRING FILLET
23 VAC
(SDS 33-42) (MPP 33-42-01)
B
RH TAXI LIGHT
Chapter 33-40
B
Page 10
170/190 MAINTENANCE TRAINING MANUAL
The navigation lights (nav lights) Three navigation lights (nav lights) assemblies are installed inside a transparent cover assembly on each wing tips front and rear side. They are turned on via the external lights panel. Each light assembly has two lamps. Normally only one lamp is on while the second lamp is on stand-by. It operates from a separate electrical source, and is activated automatically if the primary system fails. In such an event, no pilot action is necessary since a message is automatically recorded in the central maintenance computer or CMC. Located in the wing tips, main and stand-by transformers convert 115 V AC into 8.3 VAC for the green and red nav lights, and 7VAC for the white nav lights. In accordance with international law, the left nav light is red while the right one is green. Other nav lights, which are visible from behind the aircraft, are white.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 11
Figure 6: Nav lights
Right
8.3 V AC
last update: Dec06
Left
8.3 V AC
White
7 V AC
Each light assembly has two lamps. Normally only one lamp is on while the second lamp is on standby. It operates from a separate electrical source, and is activated automatically if the primary system fails.
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 12
170/190 MAINTENANCE TRAINING MANUAL
Navigation Light Transformer There are four autotransformers to supply 115VAC for operating the lamps when the aircraft is powered down. two transformers for the navigation lights are installed in the outboard portionof each wing main box. One transformer in each wing provides power for the main position lighting system and one transformer, mounted in each wing, provides power for the standby lighting system.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 13
Figure 7: Navigation Light Block Diagram
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 14
170/190 MAINTENANCE TRAINING MANUAL
The logo lights and inspection lights Logotype light switch - There is a LOGO switch installed on the cockpit overhead external lights panel which controls the logotype lights Logo type Lights - There are two lights with 75 watt sealed beam lamps. they are installed, one on top of each side of the horizontal stabilizer, and the light beam is directed to the vertical stabilizer.
Operation The logo lights are controlled by the LOGO switch, on the cockpit overhead EXTERNAL LIGHTS control panel. The logo lights interface with SPDA (Secondary Power Distribution Assembly) 2 located in the middle avionics compartment to receive 28 VDC (Volt Direct Current) power.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 15
Figure 8: Logotype Light
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 16
170/190 MAINTENANCE TRAINING MANUAL
INSPECTION LIGHT Introduction The inspection lights system provides illumination for inspection of the wing and engines by the pilot and copilot while the aircraft is flying at night or during IFR (Instrument Flight Reference) operation.
General Description There are two lamps installed in the fuselage, one on each side of the aircraft. The light beams are directed to the wing leading edge and engine intake nacelle.
Operation The inspection lights are controlled by the INSP switch, on the cockpit overhead external lights panel. The inspection lights interface with SPDA (Secondary Power Distribution Assembly) 2 located in the middle avionics compartment to receive 28 VDC (Volt Direct Current) power.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 17
Figure 9: Inspection light - block diagram
E XT ER NAL LIGHTS NAV
S TR OB E
R ED B CN
ON
ON
OFF
OFF
LOG O
T AX I NOS E
INS P S IDE
ON
ON
ON
OFF
OF F
OFF
A
LANDING LEF T
NOS E
R IG HT
ON
ON
OFF
OFF
A
CENTER FUS
I
B RH INSPECTIO N LIGHT
OVERHEAD P ANEL EXTERNAL IGHT L PANEL
A
(SDS 33-44) (MPP 33-4401)
INSPECTIONLIGHT SW INSPECTIONLTS
INSPECTION LTS
MID AVIONIC S COMPT CENTER FUS
(SDS 33-40) (MPP 33-40 -01)
I
DC POWER MODULE DC BUS 1
B
REAR PLATE RH INSPECTIO N LIGHT SPDA 2 (SSM 24-6180)
last update: Dec06
A
(SDS 33-44) (MPP 33-4401)
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 18
170/190 MAINTENANCE TRAINING MANUAL
The anti collision light system and red beacon lights The anti collision light system includes strobe lights and red beacon lights. The four strobe lights supply reference from one aircraft to another when in flight. The white 400 candle strobe-light lamps are installed inside a transparent cover assembly in the forward and rear edges of both wing tips. Associated with each strobe light is a power supply, which emits a flash rate of 50 ± 5 flashes per minute. The strobe lights are controlled by the strobe light switch located on the overhead panel. Red beacon lights are used for ground operations, to warn other traffic and airport personnel, or as a backup for the white strobe lights. The two red beacon lights are installed on top of and below the main fuselage. The dim mode, with 100 candle power, provides sufficient visibility for ground operation, while in the air the bright mode, with 400 candle power, is appropriate. The power supply for the strobe lights is inside the main fuselage, and can be replaced by maintenance within 10 to 15 minutes.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 19
Figure 10: The anti collision light system and red beacon lights
Top beacon
Red becon lights Strobe lights
The four strobe lights supply reference from one aircraft to another when in flight. The white 400 candle strobe-light lamps are installed inside a transparent cover assembly in the forward and rear edges of both wingtips. Bottom beacon last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 20
170/190 MAINTENANCE TRAINING MANUAL
Red beacon power supply There are two red-beacon power supplies. One is installed on the fuselage, adjacent to the forward cargo door, supplying power to operate the upper red beacon light. The other power supply is installed in the wing to fuselage fairing. It supplies energy to operate the lower red beacon light.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 21
Figure 11: Red Beacon Schematic
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-40
Page 22
33-50 Emergency Lights (EMB 170) Introduction The Emergency Lights System provides lighting in case the main lighting system becomes unavailable. It provides enough cabin and exterior lighting to assure safe crew and passenger evacuation even in poor visibility conditions.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 1
Figure 1: The Emergency Lights System
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 2
Emergency Light Power Unit (ELPU) The Emergency lighting system is powered by four Emergency Light Power Units (ELPUs). Two ELPUs are rigidly installed in the forward section, and two in the rear section of the fuselage to ensure power supply in case the cabin breaks apart after a crash landing. An ELPU is made up of a holder assembly, a battery pack and an electronics package. The ELPUs will automatically provide 6VDC to the emergency lights in case of a power loss on the Essential bus, if commanded by the crew or if in test mode, or activated from the flight attendant panel. The ELPUs enable emergency illumination for at least 12 minutes. The ELPUs are recharged by the 28V main battery bus.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 3
Figure 2: Emergency Light Power Unit (ELPU)
EMERGENCY LIGHT
ON / ARMED
TEST
The ELPUs will automatically provide 6VDC to the emergency lights in case of a power loss on the Essential bus, if commanded by the crew or if in test mode, or activated from the flight attendant panel.
Emergency Light Power Unit (ELP) 6 VDC
The ELPUs enable emergency illumination for at least 12 minutes.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 4
Emergency markers and identifiers All four emergency exits are marked with exit locators, markers and identifiers which are clearly visible when energized under conditions of complete darkness. The exit signs contain red letters on a white background. For general cabin emergency illumination seven floodlight assemblies are installed on the aisle ceiling panels, distributed along the fuselage. Four emergency exit area floodlight assemblies are installed at each exit. Their purpose is to illuminate the passageway leading from the main aisle to each of the four exit openings. The floor proximity emergency escape path markings are a photo luminescent type, and guide passengers to the nearest exit in conditions of dense smoke. Once outside, the passengers are provided with LED lighting on the sides of the emergency slides.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 5
Figure 3: Emergency markers and identifiers
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 6
Emergency lights system control panels The Emergency Lights System may be commanded by the emergency light switch located on the overhead panel, or by the attendant emergency light switch located on the attendant control panel installed in the forward entry area. The emergency light switch in the cockpit has three positions: • In the off position, the emergency lights are permanently turned off. This position is used before the aircraft normal electrical power or the ground power is removed. This position prevents the emergency lights from illuminating and the batteries from being drained after normal power shutdown. • In the ARM position, the emergency lights are in the stand-by mode and the batteries are charged. When normal aircraft power is lost, the emergency lights will automatically illuminate by power from the ELPU battery packs. • In the ON position, the emergency lights are manually turned on and supplied by power from the ELPU battery packs.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 7
Figure 4: The emergency control panels
The Emergency Lights System may be commanded by the emergency light switch located on the overhead panel, or by the attendant emergency light switch located on the attendant control panel.
TEMPERATURE SETTING
LOW OFF
C
HIGH
DOOR ZONE TEMPERATURE
H
CABIN LIGHTING
ON
ON
CEILING
BRIGHT DIM
SIDEWALL
BRIGHT DIM
TEST
GALLEY MASTER
ON
ON
OFF
FWD ENTRANCE
FWD GALLEY AREA
PANEL LIGHTS
BRIGHT DIM
BRIGHT DIM
TEST
COURTESY LIGHT
EMERGENCY LIGHT
ON / ARMED
ENABLED
CABIN TEMPERATURE
RESET
EVAC HORN
OFF
ON
AUTO
PSU
LAVATORY SMOKE TEST
FWD
AFT
TEST
RESET
ATTND CALL
RESET
OFF
In the off position, the emergency lights are permanently turned off. This position is used before the aircraft normal electrical power or the ground power is removed. ELPU BATTERY PACKS
OFF last update: Dec06
OFF
FOR TRAINING ONLY - Reproduction Prohibited
OFF
OFF Chapter 33-50
Page 8
Emergency light power unit The emergency light power unit is composed of sealed battery packs that provide power directly to the emergency lighting system in case of loss of primary power or when commanded by the flight crew.Each unit consists of an electronic package, plus a NiCd battery and a chassis to secure all these components.Each unit is rigidly attached to a sheet metal bracket installed between fuselage frames.The units are designed to supply power to the lights for at least 10 minutes under critical conditions.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 9
Figure 5: Emergency lights
1
MONUMENT ( REF.)
PASSENGER SIGNS EMER LT ARMED OFF
ATTND CALL
ON
STERILE
NO SMKG
C
ON
OFF
B
H
EMERGENCY LIGHT
ON/ ARMED
TEST
M MAU 2 (SSM 31-41-80)
B
GENERIC I/O MODULE COCKPIT CEILING PNL
ARM OFF
STATUS 2
FWD FUSELAGE
PASSAGEWAY LIGHT EXIT MAKER
COCKPIT EMERG LIGHT
EML4
EML4 6V
CONTROL
FWD COCKPIT CEILING PNL
FWD LH GALLERY AREA CEILING PNL
28 VDC
OVERHEAD PANEL
STATUS 1
28 VDC IN
ARM IN
OFF IN
G
J
K
MAU 1 (SSM 31-41-80)
7,5
PASS SIGN PANEL (SDS 33-23) (MPP 33-23-07)
L
C
FWD AVIONICS COMPT
GENERIC I/O MODULE
DC BUS 1
STATUS 2
CBP
STATUS 1
LH
EMER FWD BATT HTR
G
F
FSTN BELTS
ON
OFF
A COCKPIT
E
D
EXIT (SDS 33-50) (MPP 33-50-03)
D
FWD SVC DOOR LINER
EXIT LOCATOR
EXIT INDENTIFIER
EXIT
F
(SDS 33-50) (MPP 33-50-05)
EXIT
M
(SDS 33-50) (MPP 33-50-19)
E
(SDS 33-50) (MPP 33-50-15)
ARM ON OFF TEST
(SSM 33-50-80)
FWD AV COMPT
SPDA 1 (SSM 24-61-80)
EML 3
EML 5
G
CEILING PNL 1
FWD L ENTRANCE CEILING PNL EMER CABIN LIGHT
DISCRETE I/O MODULE
EMERGENCY NORMAL ON SW
EMERGENCY CABIN LIGHT
TEST SW
EMERGENCY ON INDICATOR
FWD FLIGHT ATT PANEL (SDS 33-50) (MPP 25-25-01)
EML 2
(SDS-33-50) (MPP 33-50-01)
A 1 MONUMENT (REF.)
EML 1
EMERGENCY LIGHT POWER UNIT 1
FAP ON
EXIT MAKER
FWD MAIN DOOR LINER
CEILING PNL 3 PASSAGEWAY LIGHT
EMER CABIN LIGHT
EXIT INDENTIFIER
EXIT
EXIT
FWD FUSELAGE 28 VDC
(SDS 33-50) (MPP 33-50-13)
J
(SDS 33-50) (MPP 33-50-17)
K
(SDS 33-50) (MPP 33-50-13)
H
(SDS 33-50) (MPP 33-50-15)
L
CONTROL
C
ARM
OFF
EMERGENCY LIGHT POWER UNIT 2
TEST
(SDS-33-50) (MPP 33-50-01)
EML 1
EML 2
EML 3
FAP ON
EML 5
REAR PLATE
last update: Dec06
G
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 10
Cabin attendant control panel The emergency light switch on the cabin attendant panel has two positions: • In the normal position the emergency lights remain in the mode determined by the cockpit switch. This is the normal flight position. • In the ON position the emergency lights are turned on using power from the battery packs, regardless of the cockpit emergency light switch position. The legend "ON" located in the attendant control panel is illuminated to indicate the switch selection mode. Note that the message “emergency lights not armed” will be displayed on the EICAS, and the MASTER CAUTION lights come on when the switches are selected to the ON or OFF position.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 11
Figure 6: Cabin attendant panel
CABIN LIGHTING
ON
ON
CEILING
BRIGHT DIM
SIDEWALL
BRIGHT DIM
ON
TEST
ON
FWD GALLEY AREA
PANEL LIGHTS
BRIGHT DIM
BRIGHT DIM
TEST
COURTESY LIGHT
RESET
EVAC HORN
OFF
ON
AUTO
LAVATORY SMOKE TEST
FWD
OFF
FWD ENTRANCE
EMERGENCY LIGHT
ON / ARMED
GALLEY MASTER
In the ON position the emergency lights are turned on using power from the battery packs, regardless of the cockpi emergency light switch position. The legend "ON" located in the attendant control panel is illuminated to indicate the switch selection mode.
AFT
PSU
TEST
RESET
ATTND CALL
RESET
EMER LT NOT ARMD END
ELPU BATTERY PACKS last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 12
Operation The operation of the emergency lighting system is directly related to a hazardous situation or a major failure of the normal lighting system caused by loss of the normal power.The emergency lights can be ARMED or turned ON or OFF by a switch installed in the cockpit.The exception is the photoluminescent indicator strip light that does not need to be ARMED or turned ON.Additional switches on the forward and aft flight attendant panels, allow the flight attendants to control the emergency lighting system.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 13
Figure 7: Emergency lights
C
D
E
F
G
H
B J EMERGENCY LIGHT
ON/ ARMED
TEST
A
A MID AVIONICS COMPT
L
B
K
G
RICC AFT MAIN CEILING PNL
FWD FUSELAGE
AFT L MAIN CEILING PNL
CEILING PNL 7
AFT MAIN DOOR LINER
DC BUS 2 EXIT LOCATOR
AFT LIGHT EMER BATT HTR
PASSAGEWAY LIGHT
EXIT MAKER
7.5 EXIT
EMERGENCY LIGHT POWER UNIT 3
(SSM 33-50-80)
CONTROL ARM
EXIT INDENTIFIER
EXIT
J
(SDS 33-50) (MPP 33-50-19)
28V
EMERG CABIN LIGHT
EXIT
(SDS 33-50) (MPP 33-50-09)
K
(SDS 33-50) (MPP 33-50-03)
D
L
(SDS 33-50) (MPP 33-50-15)
EML 1
EML 2
(SDS 33-50) (MPP 33-50-01)
OFF
EML 3
AFT LAVATORY TEST
AFT FLIGHT ATT PANEL (SDS 33-50) (MPP 25-25-01)
EML 5
FAP ON EML4
MAU 3
STATUS 1 STATUS 2 STATUS 2 STATUS 1
EMERGENCY NORMAL ON SW
EMERGENCY ON INDICATOR
TEST SW
G
GENERIC I/O MODULE
FWD FUSELAGE EML 4
28V
B
AFT R MAIN CEILING PNL
CEILING PNL 5
CONTROL
EMERGENCY CABIN LIGHT
ARM
EXIT MAKER
EMER CABIN LIGHT
AFT SERVICE DOOR
EXIT INDENTIFIER
EMER CABIN LIGHT
EXIT
OFF EMERGENCY LIGHT POWER UNIT 4
(SDS 33-50) (MPP 33-50-13)
(SDS 33-50) (MPP 33-50-01)
MID AVIONICS COMPT
CEILING PNL 9
C
EXIT
(SDS 33-50) (MPP 33-50-17)
H
(SDS 33-50) (MPP 33-50-13)
E
(SDS 33-50) (MPP 33-50-15)
F
EML 1
SPDA 2 (SSM 24-61-80)
EML 2 DISCRETE I/O MODUL
EML 3
TEST
EML 5
FAP ON
REAR PLATE
last update: Dec06
G
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 14
Flashlights Two flashlights are provided to help the crew during an emergency condition. The flashlight can be activated manually when it is removed from the retention bracket through a slide switch located on the flashlight body. It provides up to 45 minutes of illumination. To turn the flashlight off, you return the slide switch to the normal position. Each flashlight has a 6VDC, Ni-Cad battery which is recharged when the flashlight is inserted into its retention bracket. An internal circuit controls the battery recharging process, which can be monitored via an LED indicator, near the head of the flashlight. WARNING: Do not re-insert the flashlight with the slide switch left in the on position. This will result in overheating of and damage to the flashlight assembly.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 15
Figure 8: Flashlights position
TEMPERATURE SETTING
LOW
OFF
C
HIGH
D O O R Z O N E T E M P E R AT U R E
H
ENABLED
CABI N T EMPE RAT URE
TEMPERATURE SETTING CABIN LIGHTING
GALLEY MASTER LOW
ON
ON
ON
ON
OFF OFF
CE ILIN G
BRIGHT DIM
S ID E WALL
BRIGHT DIM
FWD E NTRANCE
TEST
FW D GALLEY AREA
BRIGHT DIM
EMERGENCY LIGHT
ON / ARMED
C
HIGH
D O O R Z O N E T E M P E R AT U R E
TEST
H
ENABLED
CABI N TEM PERAT URE
CABIN LIGHTING
EVAC HO RN
OFF
ON
ON
CE ILING
ON
S IDE WAL L
GALLEY MASTER
OFF
ON AF T E NTRANCE
PANEL LIGHTS
AUTO
LAVATORY SMOKE TEST
FWD
PANEL LIGHTS
BRIGHT DIM
COURTESY LIGHT
RESET
AFT
BRIGHT DIM
PSU
TEST
BRIGHT DIM
BRIGHT DIM
EMERGENCY LIGHT
RESET
ON / ARMED
ATTND CALL
TEST
TEST
CO UR TES Y L IG HT
RESET
OFF
E VA C H O R N
ON
AUTO
RESET
WASTE S Y STE M
FWD
TANK FULL
AFT
SERVICE TANK
FAULT
LAVATORY FAULT
ATTND CALL
WATER SY S TE M WAT ER QUANTITY
RESET
FAULT
1/4
1/2
3/4
WARNING: Do not reinsert the flashlight with the slide switch left in the on position. This will result in overheating of and damage to the flashlight assembly.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 16
190 33-50 Emergency Lights (EMB 190) Introduction The Emergency Lights System provides lighting in case the main lighting system becomes unavailable. It provides enough cabin and exterior lighting to assure safe crew and passenger evacuation even in poor visibility conditions.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 1
Figure 1: The Emergency Lights System
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 2
190 Emergency Light Power Unit (ELPU) The Emergency Lighting System is operated by 6 ELPUs, each of which supply 6VDC power.An ELPU has the capacity to allow the entire emergency lights to operate and maintain the required level of illumination for a minimum duration of 10 minutes under critical ambient conditions. It is charged and controlled by a 28VDC power source.Each ELPU has heating and quick charge features.The heating feature ensures reliable battery operation at low ambient temperatures.The quick charge feature allows a 45-minute recharge in the event the ELPU has a low charge.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 3
Figure 2: Emergency Light Power Unit (ELPU)
EMERGENCY LIGHT
ON / ARMED
TEST
The ELPUs will automatically provide 6VDC to the emergency lights in case of a power loss on the Essential bus, if commanded by the crew or if in test mode, or activated from the flight attendant panel.
Emergency Light Power Unit (ELP) 6 VDC
The ELPUs enable emergency illumination for at least 12 minutes.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 4
190 Emergency markers and identifiers All six emergency exits are marked with exit locators, markers and identifiers which are clearly visible when energized under conditions of complete darkness. The exit signs contain red letters on a white background. For general cabin emergency illumination floodlight assemblies are installed on the aisle ceiling panels, distributed along the fuselage. Emergency exit area floodlight assemblies are installed at each exit. Their purpose is to illuminate the passageway leading from the main aisle to each of the exit openings. The floor proximity emergency escape path markings are a photo luminescent type, and guide passengers to the nearest exit in conditions of dense smoke. Once outside, the passengers are provided with LED lighting on the sides of the emergency slides. The EMB 190 overwing escape path is illuminated by 6 exterior emergency lights (3 each side).
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 5
Figure 3: Emergency markers and identifiers
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 6
190 Emergency lights system control panels The Emergency Lights System may be commanded by the emergency light switch located on the overhead panel, or by the attendant emergency light switch located on the attendant control panel installed in the forward entry area. The emergency light switch in the cockpit has three positions: • In the off position, the emergency lights are permanently turned off. This position is used before the aircraft normal electrical power or the ground power is removed. This position prevents the emergency lights from illuminating and the batteries from being drained after normal power shutdown. • In the ARM position, the emergency lights are in the stand-by mode and the batteries are charged. When normal aircraft power is lost, the emergency lights will automatically illuminate by power from the ELPU battery packs. • In the ON position, the emergency lights are manually turned on and supplied by power from the ELPU battery packs.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 7
Figure 4: The emergency control panels
The Emergency Lights System may be commanded by the emergency light switch located on the overhead panel, or by the attendant emergency light switch located on the attendant control panel.
TEMPERATURE SETTING
LOW OFF
C
HIGH
DOOR ZONE TEMPERATURE
H
CABIN LIGHTING
ON
ON
CEILING
BRIGHT DIM
SIDEWALL
BRIGHT DIM
TEST
GALLEY MASTER
ON
ON
OFF
FWD ENTRANCE
FWD GALLEY AREA
PANEL LIGHTS
BRIGHT DIM
BRIGHT DIM
TEST
COURTESY LIGHT
EMERGENCY LIGHT
ON / ARMED
ENABLED
CABIN TEMPERATURE
RESET
EVAC HORN
OFF
ON
AUTO
PSU
LAVATORY SMOKE TEST
FWD
AFT
TEST
RESET
ATTND CALL
RESET
OFF
In the off position, the emergency lights are permanently turned off. This position is used before the aircraft normal electrical power or the ground power is removed. ELPU BATTERY PACKS
OFF
last update: Dec06
OFF
OFF
FOR TRAINING ONLY - Reproduction Prohibited
OFF
OFF
Chapter 33-50
OFF
Page 8
190 Emergency light power unit The emergency light power unit is composed of sealed battery packs that provide power directly to the emergency lighting system in case of loss of primary power or when commanded by the flight crew.Each unit consists of an electronic package, plus a NiCd battery and a chassis to secure all these components.Each unit is rigidly attached to a sheet metal bracket installed between fuselage frames.The units are designed to supply power to the lights for at least 10 minutes under critical conditions.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 9
Figure 5: Emergency lights 1
MONUMENT ( REF.)
PASSENGER SIGNS EMER LT ARMED OFF
ATTND CALL
ON
STERILE
NO SMKG
C
ON
OFF
OFF
B
H
EMERGENCY LIGHT
TEST
A
M MAU 2 (SSM 31-41-80)
B
GENERIC I/O MODULE COCKPIT CEILING PNL
ARM OFF
STATUS 2
FWD FUSELAGE
PASSAGEWAY LIGHT EXIT MAKER
COCKPIT EMERG LIGHT
EML4
EML4 6V
CONTROL
FWD COCKPIT CEILING PNL
FWD LH GALLERY AREA CEILING PNL
28 VDC
OVERHEAD PANEL
STATUS 1
28 VDC IN
ARM IN
OFF IN
G
J
K
MAU 1 (SSM 31-41-80)
7,5
PASS SIGN PANEL (SDS 33-23) (MPP 33-23-07)
L
C
FWD AVIONICS COMPT
GENERIC I/O MODULE
DC BUS 1
STATUS 2
CBP
STATUS 1
LH
EMER FWD BATT HTR
G
F
FSTN BELTS
ON
ON/ ARMED
COCKPIT
E
D
EXIT LOCATOR
EXIT (SDS 33-50) (MPP 33-50-03)
D
FWD SVC DOOR LINER
EXIT INDENTIFIER
EXIT
F
(SDS 33-50) (MPP 33-50-05)
EXIT
M
(SDS 33-50) (MPP 33-50-19)
E
(SDS 33-50) (MPP 33-50-15)
ARM ON OFF TEST
(SSM 33-50-80) EMERGENCY NORMAL ON SW
FWD AV COMPT
SPDA 1 (SSM 24-61-80)
EML 3
EML 5
G
CEILING PNL 1
FWD L ENTRANCE CEILING PNL EMER CABIN LIGHT
EMERGENCY CABIN LIGHT
TEST SW
EMERGENCY ON INDICATOR
FWD FLIGHT ATT PANEL (SDS 33-50) (MPP 25-25-01)
EML 2
(SDS-33-50) (MPP 33-50-01)
A 1 MONUMENT (REF.)
EML 1
EMERGENCY LIGHT POWER UNIT 1
FAP ON
DISCRETE I/O MODULE
EXIT MAKER
CEILING PNL 3 PASSAGEWAY LIGHT
FWD MAIN DOOR LINER
EMER CABIN LIGHT
EXIT INDENTIFIER
EXIT
EXIT
FWD FUSELAGE 28 VDC
(SDS 33-50) (MPP 33-50-13)
J
(SDS 33-50) (MPP 33-50-17)
K
(SDS 33-50) (MPP 33-50-13)
H
(SDS 33-50) (MPP 33-50-15)
L
CONTROL
C
ARM
OFF
EMERGENCY LIGHT POWER UNIT 2
TEST
(SDS-33-50) (MPP 33-50-01)
EML 1
EML 2
EML 3
FAP ON
EML 5
REAR PLATE
last update: Dec06
G
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 10
190 NOTES:
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 11
Figure 6: Emergency lightning - block diagram E
P AS S ENGE R SIGNS E ME R LT ARMED OFF
S TE RILE
D
AT T ND C ALL
ON
NO SMKG
F ST N BE LT S
ON
ON
OFF
OF F
B ON/ ARMED
H
T ES T
A
H G
C
MONUMENT ( RE F.)
B MID AVIONICS COMPT
COCKPIT
LICC
DC BUS
F
D
C 1
J
F WD C EILING (RE F.)
E ME RG ENCY LIG HT
LH CBP DC BUS
1
EMER LT BATT HTR
1
OVERHEAD PANEL
MAU 2 (SSM 31-41-80)
PASSENGER SIGN PANEL (SDS 33-23) (MPP 33-23-07)
A
1
MID AVIONICS COMPT
FWD AVIONICS COMPT
MAU 3 (SSM 31-41-80)
BACKPLANE
FWD FAP (SSM 25-25-01)
BACKPLANE
G
MONUMENT (R EF .)
C
EMERGENCY ON INDICATOR
EMER FWD BATT HTR
ARM IN OFF IN
ARM ON OFF
FWD AVIONICS COMPT GENERIC I/O MOD. 9
GENERIC I/O MOD. 9
SPDA 1 (SSM 24-61-80) TEST SW
28 VDC
28 VDC
D
F
REAR PLACE
K
DISCRETE I/O (SLOT 11)
(TY PICAL)
6 VDC
28 VDC
EMERGENCY NORMAL ON SW
OVE RW ING E XIT TEST
2 STATUS
FAP ON
1 STATUS
EMERGENCY LIGHT POWER UNIT 3
FWD RH EXT EMERG LT
RH EMERG DOOR LINER
CABIN EMERG LT
EXIT LOCATOR
EXIT MARKER
(SDS 33-50) (MPP 33-50-19)
last update: Dec06
(SDS 33-50) (MPP 33-50-13)
G
(SDS 33-50) (MPP 33-50-17)
6 VDC
EXIT MARKER
H
(SSM 33-50-80 SHEET 4)
L
E
EXIT IDENTIFIER
EXIT
EMERGENCY LIGHT POWER UNIT 4
EXIT
J
F
RH EMERG DOOR LINER
LH EMERG DOOR LINER
MIDDLE CEILING PANEL
EXIT
K
EML 5
EML 3
CABIN CEILING PANEL
MID OVERWING CEILING PANEL
EML 3
6 VDC
6 VDC
6 VDC
EML 2
EML 1
(SDS-33-50) (MPP 33-50-01)
FAIRING
(SDS 33-50) (MPP 33-50-25)
EML 4
OFF
CONTROL
CHARGE/HEAT
ARM
MIDDLE CEILING PANEL
(SDS 33-50) (MPP 33-50-17)
FOR TRAINING ONLY - Reproduction Prohibited
EXIT
H
(SDS 33-50) (MPP 33-50-15)
L
1
A MONUMENT CAN BE A LAVATORY OR A CLOS ET
Chapter 33-50
Page 12
190 NOTES:
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 13
Figure 7: Emergency lightning - block diagram (continued) F WD C EILING (RE F.)
G A
B
E ME RG ENCY LIG HT
ON/ ARMED
T ES T
D
E
H
H
J
P AS S ENGE R SIGNS E ME R LT ARMED OFF
B
ON
S TE RILE
NO SMKG
MID AVIONICS COMPT
COCKPIT
LICC
DC BUS
ON
ON
OF F
EMER LT BATT HTR
C LH CBP
DC BUS
1
F ST N BE LT S
OFF
A
1
OVERHEAD PA NEL
G
MAU 2 (SSM -41-80) 31
C
AFT LH LAVAT ORY
MID AVIONICS COMPT
FWD AVIONICSCOMPT
PASSENGER GN SI PANEL (SDS 33-23 ) (MPP 33-23 -07)
(TY PICAL)
D
AT T ND C ALL
MAU 3 (SSM1-41-80) 3
BACKPLANE
MID AVIONICS COMPT
AFT FAP (SSM 25-25-83)
SPDA 2 (SSM 24-61-8 0)
B
BACKPLANE
REAR PLACE
EMERGENCY ON INDICATOR
EMER FWD BATT HTR
K
ARM IN OFF IN
DISCRETE I/O (SLOT 9)
GENERIC I/O MOD. 9
GENERIC I/O MOD. 9
ARM
TEST WS
ON
EMERGENCY NORMAL ON SW
28 VDC
28 VDC
28 VDC
OFF
(TY PICAL)
6 VDC
OVE RW ING E XIT
RH EMERG DOOR LINER
MID RH EXT EMERG LT
J
last update: Dec06
K
(SDS 33 -50) (MPP 33-50-17)
H
(SSM 33-50-80 SHEET 3)
TEST
2 STATUS
FAP ON
1 STATUS
5 EML
E
EXIT IDENTI FIER
EXIT
EMERGENCY LIGHT POWER UNIT 3
EXIT
(SDS 33-50) (MPP 33-50-25)
EXIT MARKER
EML
EXIT MARKER
F
LH EMERG DOOR LINER
LH EMERG DOOR LINER
MIDDLE CEILING PANEL
3
FAIRING
FWD LH EXT EMERG LT
G
6 VDC
3 EML 6 VDC
2 EML 6 VDC
1 6 VDC
EML
(SDS-33-50) (MPP 33-50-01)
FAIRING
(SDS 33 -50) (MPP 33-50-25)
4 EMERGENCY I LGHT POWER UNIT 4
EML
OFF
CONTROL
CHARGE/HEAT
ARM
MIDDLE CEILING EL PAN
(SDS 33 -50) (MPP 33-50-17)
FOR TRAINING ONLY - Reproduction Prohibited
EXIT
H
(SDS 33 -50) (MPP 33-50-15)
F
Chapter 33-50
Page 14
190 Cabin attendant control panel The emergency light switch on the cabin attendant panel has two positions: • In the normal position the emergency lights remain in the mode determined by the cockpit switch. This is the normal flight position. • In the ON position the emergency lights are turned on using power from the battery packs, regardless of the cockpit emergency light switch position. The legend "ON" located in the attendant control panel is illuminated to indicate the switch selection mode. Note that the message “emergency lights not armed” will be displayed on the EICAS, and the MASTER CAUTION lights come on when the switches are selected to the ON or OFF position.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 15
Figure 8: Cabin attendant panel
CABIN LIGHTING
ON
ON
CEILING
BRIGHT DIM
SIDEWALL
BRIGHT DIM
ON
TEST
ON
FWD GALLEY AREA
PANEL LIGHTS
BRIGHT DIM
BRIGHT DIM
TEST
COURTESY LIGHT
RESET
EVAC HORN
OFF
ON
AUTO
LAVATORY SMOKE TEST
FWD
OFF
FWD ENTRANCE
EMERGENCY LIGHT
ON / ARMED
GALLEY MASTER
In the ON position the emergency lights are turned on using power from the battery packs, regardless of the cockpi emergency light switch position. The legend "ON" located in the attendant control panel is illuminated to indicate the switch selection mode.
AFT
PSU
TEST
RESET
ATTND CALL
RESET
EMER LT NOT ARMD END
ELPU BATTERY PACKS
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 16
190 Operation The operation of the emergency lighting system is directly related to a hazardous situation or a major failure of the normal lighting system caused by loss of the normal power.The emergency lights can be ARMED or turned ON or OFF by a switch installed in the cockpit.The exception is the photoluminescent indicator strip light that does not need to be ARMED or turned ON.Additional switches on the forward and aft flight attendant panels, allow the flight attendants to control the emergency lighting system.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 17
Figure 9: Emergency lights C
D
E
F
G
H
B J EMERGENCY LIGHT
ON/ ARMED
TEST
A
A MID AVIONICS COMPT
L
B
K
G
RICC AFT MAIN CEILING PNL
FWD FUSELAGE
AFT L MAIN CEILING PNL
CEILING PNL 7
AFT MAIN DOOR LINER
DC BUS 2 EXIT LOCATOR
AFT LIGHT EMER BATT HTR
PASSAGEWAY LIGHT
EXIT MAKER
7.5 EXIT
EMERGENCY LIGHT POWER UNIT 3
CONTROL
(SSM 33-50-80)
J
(SDS 33-50) (MPP 33-50-19)
28V
ARM
EMERG CABIN LIGHT
EXIT INDENTIFIER
EXIT
EXIT
(SDS 33-50) (MPP 33-50-09)
K
D
(SDS 33-50) (MPP 33-50-03)
L
(SDS 33-50) (MPP 33-50-15)
EML 1
EML 2
(SDS 33-50) (MPP 33-50-01)
OFF
EML 3
AFT LAVATORY TEST
AFT FLIGHT ATT PANEL (SDS 33-50) (MPP 25-25-01)
EML 5
FAP ON EML4
MAU 3
STATUS 1 STATUS 2 STATUS 2 STATUS 1
EMERGENCY NORMAL ON SW
TEST SW
EMERGENCY ON INDICATOR
G
GENERIC I/O MODULE
FWD FUSELAGE EML 4
28V
B
AFT R MAIN CEILING PNL
CEILING PNL 5
CONTROL
EMERGENCY CABIN LIGHT
ARM
EXIT MAKER
AFT SERVICE DOOR
EXIT INDENTIFIER
EMER CABIN LIGHT
EXIT
OFF EMERGENCY LIGHT POWER UNIT 4
(SDS 33-50) (MPP 33-50-13)
(SDS 33-50) (MPP 33-50-01)
MID AVIONICS COMPT
CEILING PNL 9
EMER CABIN LIGHT
C
EXIT
(SDS 33-50) (MPP 33-50-17)
H
(SDS 33-50) (MPP 33-50-13)
E
F
(SDS 33-50) (MPP 33-50-15)
EML 1
SPDA 2 (SSM 24-61-80)
EML 2 DISCRETE I/O MODUL
EML 3
TEST
EML 5
FAP ON
REAR PLATE
last update: Dec06
G
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 18
190 Flashlights Two flashlights are provided to help the crew during an emergency condition. The flashlight can be activated manually when it is removed from the retention bracket through a slide switch located on the flashlight body. It provides up to 45 minutes of illumination. To turn the flashlight off, you return the slide switch to the normal position. Each flashlight has a 6VDC, Ni-Cad battery which is recharged when the flashlight is inserted into its retention bracket. An internal circuit controls the battery recharging process, which can be monitored via an LED indicator, near the head of the flashlight. WARNING: Do not re-insert the flashlight with the slide switch left in the on position. This will result in overheating of and damage to the flashlight assembly.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 19
Figure 10: Flashlights position
TEMPERATURE SETTING
LOW
OFF
C
HIGH
D O O R Z O N E T E M P E R AT U R E
H
ENABLED
CABI N T EMPE RAT URE
TEMPERATURE SETTING CABIN LIGHTING
GALLEY MASTER LOW
ON
ON
ON
ON
OFF OFF
CE ILIN G
BRIGHT DIM
S ID E WALL
BRIGHT DIM
FWD E NTRANCE
TEST
FW D GALLEY AREA
BRIGHT DIM
EMERGENCY LIGHT
ON / ARMED
TEST
C
HIGH
H
ENABLED
CABI N TEM PERAT URE
CABIN LIGHTING
EVAC HO RN
OFF
ON
ON
CE ILING
ON
S IDE WAL L
GALLEY MASTER
OFF
ON AF T E NTRANCE
PANEL LIGHTS
AUTO
AFT
BRIGHT DIM
PSU
LAVATORY SMOKE TEST
FWD
D O O R Z O N E T E M P E R AT U R E
PANEL LIGHTS
BRIGHT DIM
COURTESY LIGHT
RESET
TEST
BRIGHT DIM
BRIGHT DIM
EMERGENCY LIGHT
RESET
ON / ARMED
ATTND CALL
TEST
TEST
CO UR TES Y L IG HT
RESET
OFF
E VA C H O R N
ON
AUTO
RESET
WASTE S Y STE M
FWD
TANK FULL
AFT
SERVICE TANK
FAULT
LAVATORY FAULT
ATTND CALL
WATER SY S TE M WAT ER QUANTITY
RESET
FAULT
1/4
1/2
3/4
WARNING: Do not reinsert the flashlight with the slide switch left in the on position. This will result in overheating of and damage to the flashlight assembly.
last update: Dec06
FOR TRAINING ONLY - Reproduction Prohibited
Chapter 33-50
Page 20
170/190 MAINTENANCE TRAINING MANUAL
33-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 23-00 Passenger Signs ¦ ¦ ¦ ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦
(M)(O)No passenger, flight attendant seat or lavatory may be occupied from which a "No Smoking/Fasten Seat Belt" sign is not readily legible, or that seat must be blocked and placarded "DO NOT OCCUPY".
¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
C ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
(O)"No Smoking/Fasten Seat Belt" signs may be inoperative and the affected passenger seat(s), flight attendant seat(s) or lavatory may be occupied provided: a) The passenger address system operates normally and can be clearly heard throughout the cabin during flight, and b) The passenger address system is used to notify the flight attendant and passengers when seat belts should be fastened and when smoking is prohibited.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 23-09 Cockpit Sterile ¦ Light ¦
C ¦ ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
¦ 26-00 Courtesy Lights ¦ System ¦
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ sufficient light is available at ¦ passenger entry area.
¦ ¦ ¦
¦ 31-00 Forward and Aft C ¦ 9 ¦ Cargo Compartment ¦ ¦ Lights ¦
¦ 0 ¦ ¦
¦ ¦ ¦
¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: ORIGINAL ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 12/16/2003 ¦ 33-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 33 LIGHTS ¦ ¦ ¦
¦ Individual lights may be ¦ inoperative provided remaining ¦ lights are: ¦ a) Sufficient to clearly ¦ illuminate all required ¦ instruments, controls, and ¦ other devices for which they ¦ are provided, ¦ b) Positioned so that direct ¦ rays are shielded from ¦ flight crewmembers' eyes, ¦ c) Lighting configuration and ¦ intensity is acceptable to ¦ the flight crew, and ¦ d) Flight deck emergency ¦ lighting is operative.
¦ Individual lights may be ¦ inoperative provided: ¦ a) Sufficient lighting remains ¦ for flight attendants to ¦ perform their assigned ¦ duties, ¦ b) No more than 10 per cent ¦ ceiling light lamps are ¦ inoperative, ¦ c) No more than 2 adjacent ¦ ceiling light lamps in the ¦ longitudinal or lateral ¦ direction are inoperative, ¦ d) Remaining operational ¦ ceiling, forward entry area, ¦ forward galley area and aft ¦ entry area lighting must be ¦ functional in BRIGHT ¦ setting, and ¦ e) Cabin emergency lighting is ¦ operative.
MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 31-00 Air Data Smart ¦ Probe (ADSP) ¦ Heater ¦ Controllers
B ¦ 8 ¦ ¦ ¦
¦ 4 ¦ ¦ ¦
¦ One Heater controller per ADSP may ¦ be inoperative. ¦ ¦
¦ ¦ ¦ ¦
¦ 41-00 Windshield Wiper ¦ Systems ¦ ¦ ¦
C ¦ 2 ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦
¦ ¦ ¦
1) Low Speed Mode C ¦ 2 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ associated High Speed Mode operates ¦ normally.
¦ ¦ ¦
¦ ¦ ¦
2) High Speed Mode
C ¦ 2 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ associated Low Speed Mode operates ¦ normally.
¦ ¦ ¦
¦
3) Timer Mode
C ¦ 2
¦ 0
¦
¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
4) Parking Mode
C ¦ 2 ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 0 ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ (M)May be inoperative provided: ¦ a) Blades are positioned to ¦ provide an acceptable field ¦ of vision to flight crew, ¦ and ¦ b) Associated Windshield Wiper ¦ System is considered ¦ inoperative and not used.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 42-00 Windshield ¦ Heating Systems ¦
B ¦ 2 ¦ ¦
¦ 1 ¦ ¦
¦ One may be inoperative provided ¦ airplane is not operated in known ¦ or forecast icing conditions.
¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: ORIGINAL ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 12/16/2003 ¦ 33-3 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 33 LIGHTS ¦ ¦ ¦
¦ (M)Any light may be inoperative ¦ provided the following minimum ¦ configuration is complied with: ¦ a) One green light at the right ¦ forward wing tip position, ¦ b) One red light at the left ¦ forward wing tip position, ¦ and ¦ c) One white light at each aft ¦ wing tip position.
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ ¦
C ¦ 8 ¦
¦ 0 ¦
¦ May be inoperative for day ¦ operations.
¦ ¦
¦ 44-00 Wing Inspection ¦ Lights ¦
C ¦ 2 ¦ ¦
¦ 0 ¦ ¦
¦ (O)May be inoperative provided ¦ ground deicing procedures do not ¦ require their use.
¦ ¦ ¦
¦ 45-00 Red Beacon Lights C ¦ 2 ¦ ¦
¦ 0 ¦
¦ May be inoperative provided strobe ¦ lights operate normally.
45-00 Aircraft Diagnostic and Maintenance System (Central maintenance System). Introduction The Aircraft Diagnostic and Maintenance System is a centralized interface that the mechanic uses to perform most maintenance activities on the aircraft.The maintenance system is a combination of a fault recording and maintenance access system and the LRUs/LRMs that are member systems. The ADMS is comprised of a Central Maintenance Computer (CMC) that hosts the Central Maintenance Computer Function and the Aircraft Condition Monitoring Function (ACMF), the ACMS ground support software Report Builder and Report Analyser, an optional Remote Terminal (RT), and an optional Data Management Unit (DMU).
General Description The CMC supplies the operator access to the member systems from a single user interface.The access can be from the RT (Remote Terminal) through the LAN (Local Area Network) or from the copilot’s MFD (Multi-Function Display) through the RIB (Remote Image Bus).The copilot’s access is with the CCD (Cursor Control Device). The LAN ports are found in the forward avionics compartment, cockpit, external maintenance panel and middle avionics compartment. NOTE : The term member system refers to any system installed on the aircraft that complies with the CMC interface requirements and implements the features of the CMC. The CMC is connected to a aircraft battery (HOT BATT BUS 2) for backup power purposes.
The CMC system hosts: • CMCF (Central-Maintenance-Computer Function) • ACMS (Aircraft Condition Monitoring System) ground support software
The CMC system has the components that follow: • CMC module • Database module • DMU (Data-Loader Management Unit)
Figure 1: CMC Introduction
PRINTER
ASCB MEMBER SYSTEM
ASCB MEMBER SYSTEM
AIRCRAFT BATTERY
COPILOTS MFD
CCD
28V DC ASCB
TO OTHER MAUs
REMOTE IMAGE BUS
LAN
MAU
I/O MODULE
NIC
CENTRAL MAINTENANCE COMPUTER MODULE
DATABASE MODULE
ARINC 429
BACKPLANE
429 MEMBER SYSTEM
429 MEMBER SYSTEM
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Notes:
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Components CMC MODULE The CMC module is the central controller for the CMC system. The CMC module interfaces with other elements of the system through the virtual backplane.It hosts the functional software, fault history database, loadable diagnostic information database, and ACMF (Aircraft ConditionMonitoring Function) data. The CMC module is powered by the DC (Direct Current) bus 1.The CMCF is powered by the MAU (Modular Avionics Unit) power supply through the backplane and receives its data from the ASCB (Avionics Standard-Communication Bus).If the MAU is not powered or the ASCB is not in use, the maintenance function is not available.The CMC module is also connected to the aircraft battery (HOT BATT BUS 2) for backup power purposes.The aircraft battery connection is used to power down the CMCs operating system in the event that the CMC module loses electrical power from the DC bus 1. The CMC module has circuitry that prevents the CMC module from using more than two minutes of the aircraft battery.In the event that the CMCF does not power down within two minutes, the CMC module hardware automatically disconnects the CMC module from the aircraft battery power.
Figure 2: CMC Location
CMC Location MAU 1 #
B U S
C H
20 B 19 2 B 18 2 B 17 2 B
16 2 B 15 14 2 B 13 2 B 2 B 12 2 B 11 10 9
8 7 6 5 4 3 2 1
2 B
C H
B U
#
S
CMC GPS 1 Power Supply 2 ESS 1 FCM 1 A 1
16 15 14 13 12 11 10 9
AIOPB1 PROC 1 NIC 1 (A) (ID = 1) FCM 2
AIOPA1
A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1 A 1
2 B
SPARE SPARE GENERIC I/O 2
U
C H
8 7 6 5 4 3 2 1
NIC 4 (B) (ID = 61) PROC 4 PROC 3 NIC 3 (A) (ID = 29) SPARE DATABASE AUTOBRAKE EGPWM NOSEWHEEL STEERING AGM 2
2 B 2 B
B
#
MAU 3 C H
B U
U S
Power Supply 1 DC 2
C H
S
C H
Power Supply 1 ESS 1
C H
B U S
B U S
C H
16 1 B 15 14 13 12 1 B 11 10 1 B 9 1 B 1 B
A 1 A 1
A 1 A 1 A 1 A C H
8 7 6 5 4 3 2 1
1 B
1 B U
PROC 1 = ADA 1, MW 1, UTIL 1, CAL/MCDU 1, CMS 1
#
C H
B U S
A 2 2 A
FCM 3 A 2 GENERIC I/O 3 A 2 NIC 6 (B) (ID = 30) PROC 6 PROC 5 NIC 5 (A) (ID = 33) CUSTOM I/O 2
A 2 A 2 A 2 A 2
SPARE SPARE FCM 4 A 2
B
U S
Power Supply 2 DC 2 ENGINE VIBE GPS 2 PSEM 2
AIOPB2
B
PROC 2 = CMF 2 U
#
S
A 1 2 B 2 B
A 1 NIC 2 (B) (ID = 62) PROC 2 GENERIC I/O 1
2 B 2 B 2 B
Power Supply 2 ESS 2/DC 2 BRAKES (INBD) CONTROL I/O 2 AIOPA2
Central Maintenance computer Module (CMCM). Central Maintenance computer Function (CMCF). Issue: June06 Revision: 00
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Data Management Unit (DMU) The CMC has a LAN connection to the DMU. The DMU is a combination DVD-ROM drive and two PCMCIA (Personal Computer Memory Card International Association) type II and III slots.The drives supply the user the capability to transfer files to and from the aircraft.The software loading is done with a data loading function on the CMCF. The DMU DVD-ROM is powered only while the aircraft is on the ground through a discrete from the MAU. The DMU is a LRU (Line Replaceable Unit) installed in the RH (Right-Hand) aft console.The DMU is powered by DC bus 2.
Figure 3: Data Management Unit (DMU)
DMU LOCATION
Supports: DVD ROM (CD ROM) PCMCIA CARDS.
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Remote Terminal (RT) The Remote Terminal (RT) is an optional user interface that can be connected to the aircraft at several locations.The ERJ-170 remote terminal ports are located in the cockpit, external maintenance panel, and forward as well as mid electronic compartments.This allows the mechanic access to the Aircraft Diagnostic and Maintenance System while the pilot is in the cockpit or passengers are loading or unloading from the aircraft.This also allows the mechanic to access maintenance information close to the equipment under repair. The RT is a lap top PC with a Local Area Network (LAN), Ethernet 10Base2 card, CD-ROM, and 3.5 inch, 1.44 MB floppy disk drive.In order to use the RT for ADMS functionality, the CMC module needs to be operating.This requires the Central Maintenance Computer (CMC) to be powered and the ASCB-D data bus operating. If the CMC is not operating the LAN and the RT can still be used for functionality other than ADMS.If the CMC is operating but the ASCB-D data bus is failed the FHDB can still be accessed but no other maintenance functionality is available. The remote terminal ADMS software functionality and GUI is the same as that provided in the cockpit using the copilot’s MFD and CCD.Additionally the Remote Terminal is the only display that the mechanic can use to view aircraft maintenance manuals.Note that only one user of the CMC is supported at any time.The remote terminal also allows to load software to the aircraft via a Data Loading System (DLS) function resident on the PC.
Figure 4: Remote Terminal - Ports
The EMBRAER 170 remote terminal ports are located in the cockpit, external AC power panel, and forward as well as mid electronic compartments
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Notes:
Figure 5: Local Area Network (LAN)
123 BL (REF.)
E
D
C
B A
A
125 BL (REF.)
FWD RAMP PNL
C MAIN PANEL
FWD AVIONICS COMPT
D
B COCKPIT
MAIN PANEL
COCKPIT
MAIN PANEL
COCKPIT
MAU 1 (SSM 31-41-80)
EICAS (SSM 31-61-80)
LAN PORT (SDS 45-45) CMC MODULE (SSM 45-45-80)
T
NIC 2
MFD 1 (SSM 31-61-80)
FWD AVIONICS COMPT
FWD AVIONICS COMPT
MRC 1 (SSM 34-02-80)
SPDA 1 (SSM 24-61-80)
FWD AVIONICS COMPT
E
PFD 1 (SSM 31-61-80)
NIC 1
NIM
FWD AVIONICS COMPT MAU 2 (SSM 31-41-80)
LAN PORT (SDS 45-45)
ASCB MODULE
NIC 3
NIC 4
C
F T
ASCB MODULE
NIM
NIC 6
NIC 5 DMU (SSM 45-45-80)
LAN PORT (SDS 45-45)
L MID AVIONICS COMPT
SPDA 2 (SSM 24-61-80)
MRC 2 (SSM 34-02-80)
MAU 3 (SSM 31-41-80)
MID AVIONICS COMPT
MID AVIONICS COMPT
MID AVIONICS COMPT
PFD 2 (SSM 31-61-80)
LAN PORT (SDS 45-45)
MFD 2 (SSM 31-61-80)
H RH AFT CONSOLE
COCKPIT
RH CONSOLE
MAIN PANEL
COCKPIT
MAIN PANEL
H
COCKPIT
L
G
K J
F
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G
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Central Maintenance System CENTRAL MAINTENANCE COMPUTER The CMC communicates with and receives communication from the Member Systems either directly through ASCB-D for ASCB Member Systems, or through ASCB-D via the MAU I/O and ARINC 429 buses for ARINC 429 Member Systems.The CMCM hosts the CMCF software and provides storage for the LDI, the fault history database (FHDB), ACMF data, and Windows NT operating system.The CMC Module’s Bus Interface Controller (BIC) ensures separation of the non-critical Level D Maintenance System data from critical data that exists on ASCB-D.The BIC provides both a hardware and software fire wall to ensure that the Level D CMCF software can not interfere with critical data on the ASCB-D.
The ACMS is a function that is used to extend the capabilities of the CMC system.The ACMS is made up of the ACMF and the RB (Report Builder). The ACMF is a data-driven software application that increases the operator’s ability to isolate the flight anomalies The ACMF collects a set of predefined data based on a defined trigger even that uses the ground support software RB.The ACMS supplies the function that follows: • Monitors aircraft parameters • Records aircraft parameters based on real-time evaluation of monitored aircraft parameters through the trigger logic • Supplies access to reports through the download and printout • Lets operator develop more custom monitoring capabilities for any aircraft system • Supplies storage for certain accumulated values AIRCRAFT CONDITION MONITORING FUNCTION (ACMF)
LOADABLE DIAGNOSTIC INFORMATION (LDI) The maintenance system uses the LDI (Loadable Diagnostic Information) as a data model or map that characterizes the member systems maintenance data.The maintenance data includes maintenance screens and messages. The LDI data is stored in a database that is not part of the CMCF executable code, but instead is in a separate file accessed by the CMCF.The maintenance screens data is stored in the LDI.The LDI can be updated through the data loading system. FAULT HISTORY DATABASE All fault occurrences are stored in the FHDB in the CMC module.The FHDB is limited to 10 MB total. A maintenance message record is equivalent to 57 bytes.When the storage limit occurs the oldest records are deleted first. The FHDB redundancy and aircraft fault history are kept in a copy of the FHDB in the aircraft database module. If a CMC module is replaced, the fault history of the aircraft is kept. AIRCRAFT CONDITION MONITORING SYSTEM (ACMS)
The ACMF runs on the CMC and supplies a procedure to complement the CMCF.The ACMF supplies storage and analysis of system fault data, with storage of more data that is not directly associated with faults.This stored data can be downloaded through the RB for analysis. The ACMF monitors the ASCB data to determine if conditions satisfy the trigger logic for any applications.Once a trigger condition is satisfied, data associated with the trigger condition is stored in the CMC module. Periodically the ACMF reports are stored on the database module.The stored data can be retrieved for an in-air download through the CMF or the printout on the LAN printer. The CMC module lets up to 25 megabytes of storage for the ACMF reports. The RB lets reports be categorized into storage families.The ACMF supplies a memory full flag that is set when the total available storage exceeds 90% of available space.This flag is shown in the RB when it accesses the ACMF. The notebook PC, running the RT application, uses the aircraft LAN and the RB to download reports (via the RT) and clean up storage.Once a family has reached its limit the oldest report within a family is over written as a new report is created.
Figure 6: Central Maintenance Computer
Modular Avionics Units (MAU)
ARINC Member Systems
BIC
I/O Module
(running Digital Engine Operating System)
Software Services (Core S/W, PDD) Level A
I/O Software
ARINC 429
I/O Process Level B Maintenance Process Level C
Remote Terminal
Test Panel
(Optional)
External Power Source
Virtual Backplane Bus (PCI)
CMC Module (running COTS Operating System)
Backplane Interface Controller Level A
Aircraft Fault History Central Maintenance Software Fault Processing - Level ED Aircraft Condition Monitoring - Level E I/O Processing - Level E
ACMF Data LDI Database
Time Limited Dispatch Aircraft Fault History
BIC
ACMF Data LDI Database
PROCESSOR MODULE BIC
(Optional)
Remote Image Bus (RIB) ARINC 429
DATABASE MODULE (Backup Data Files)
Data Management Unit
Copilot CCD
Copilot Multi-Function Display ASCB Member Systems
Time Limited Dispatch
NIC BIC
Ethernet LAN
ASCB-D
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REPORT BUILDER (RB) The RB is a windows-based software application that is used to set the data file used by the ACMF application.The RB is used to calculate the data parameters to monitor,calculate what combination of data parameters are necessary to start the reports, and calculate what parameters are to be recorded in a report when a given report is started.The RB runs on a ground-based computer that uses the windows NT operating system.The RB allows for the transfer of the applications, to retrieve triggered reports, and clean up the storage in the CMC module.
REMOTE TERMINAL (RT) The RT is an application that runs on a notebook PC.The RT is an user interface that can be connected to the aircraft, through the LAN ports, located in the cockpit, external maintenance panel, forward avionics compartment and middle avionics compartment. This lets the mechanic access maintenance data close to the equipment under repair while the passengers get on or off the aircraft.The RT functionality and graphic user interface is the same as that provided in the cockpit using the copilot’s MFD and CCD. By using the RT application, the mechanic will be able to access the CMC pages, generate ACMF reports and/or transfer field-loadable software to the aircraft systems.
CMC SYSTEM INTERFACE The CMC system interfaces with the components and functions that follow: • Copilot’s MFD: The copilot’s MFD is a user interface for the CMCF with control of the display supplied by the CCD.The CMCF data is only shown on the copilot’s MFD. • RIB: The CMC module has a direct connection to the copilot’s MFD through a RIB. The DU (Display Unit)s software directs the CMCFs video to the display.
• CCD: The CCD is used to control the maintenance display from the copilot’s seat. • MWF (Monitor Warning Function): The MWF continuously provides the CMCF a list of all CAS messages and status of each messages.The CMCF compares the CAS messages with the maintenance messages and stores this data in the FHDB.The monitor warning system monitors the CMC status through the ASCB. • NIC (Network Interface Controller): The NIC supplies the gateway between the ASCB and the backplane.The NIC supplies time, data, aircraft serial number, and aircraft type to the CMCF and the member systems.The NIC also supplies timimg data related to data transmissions within the system. • Member Systems: The CMC member systems communicate with the CMCF through the ASCB and the ARINC-429.The member system parameter group includes all fault and identification data related to the member system The I/O (Input/Output) module helps the CMCF to receive the ARINC-429 labels and place them onto ASCB in a member system maintenance data parameter group.The I/ O module also receive data on the ASCB from the CMCF.The CMCF maintenance data parameter group transmits to the ARINC-429 member systems.The I/O module supplies the support of the file transfer from the ARINC-429 member systems.
Figure 7: CMC Module
Modular Avionics Units (MAU)
ARINC Member Systems
BIC
I/O Module
(running Digital Engine Operating System)
Software Services (Core S/W, PDD) Level A
I/O Software
ARINC 429
I/O Process Level B Maintenance Process Level C
Remote Terminal
Test Panel
(Optional)
External Power Source
Virtual Backplane Bus (PCI)
CMC Module (running COTS Operating System)
Backplane Interface Controller Level A
Aircraft Fault History Central Maintenance Software Fault Processing - Level ED Aircraft Condition Monitoring - Level E I/O Processing - Level E
ACMF Data LDI Database
Time Limited Dispatch Aircraft Fault History
BIC
ACMF Data LDI Database
PROCESSOR MODULE BIC
(Optional)
Remote Image Bus (RIB) ARINC 429
DATABASE MODULE (Backup Data Files)
Data Management Unit
Copilot CCD
Copilot Multi-Function Display ASCB Member Systems
Time Limited Dispatch
NIC BIC
Ethernet LAN
ASCB-D
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• Ground Interlock Switch: The member system uses the ground interlock switch as an additional safety interlock when maintenance actions are done on the aircraft that require more safety.The member systems supply the interlocks so that if the software in the CMC fails, the member system cannot be commanded to any unsafe condition or start tests while the aircraft is in flight. • Printer: The CMC interfaces with the printer through the LAN and the printer interfaces to the MAU 2 through the ARINC-429.All communications with the printer from any other systems, such as the CMU (Communications Management Unit), are accomplished through the CMC.The printer supplies all print reports to the CMC system through the LAN, and all fault reports through the ARINC-429 connection.
MAINTENANCE MODE The maintenance mode is only accessible when the aircraft is on ground and safety conditions are met. The maintenance mode functions and access to the maintenance screens are locked out by both the CMCF and the member systems.This supplies interlocks so that if the CMC fails, the member system cannot be commanded to any unsafe condition or starts the tests when not in other than maintenance mode.This mode is used by the maintenance crew to find and repair the member system.When the CMCF is in the maintenance mode it supplies access for the member systems to show fault data (active faults).The command systems diagnostics start the BIT.The member system supplies defined-passive status screens (real-time display of system status) and commands the download of stored fault data (NVM download).During the maintenance mode the member system requirements are as follow: • Transmission of the member system finds fault,unless the member system is doing a commanded test or a NVM download. • Transmission of the equipment Id,SDI (Source/Destination Identifier),hardware part number,software part number,and serial number information,unless the aircraft system is doing a commanded test or a NVM download
• Performance of power-up and continuous BIT to find faults unless the member system is doing a commanded test or a NVM download. • Performance of commanded test (if commanded by the CMC,if necessary and sufficient interlocks are satisfied). • Performance of NVM download (if commanded by the CMC,if necessary and sufficient interlocks are satisfied).
Figure 8: CMC Module
Modular Avionics Units (MAU)
ARINC Member Systems
BIC
I/O Module
(running Digital Engine Operating System)
Software Services (Core S/W, PDD) Level A
I/O Software
ARINC 429
I/O Process Level B Maintenance Process Level C
Remote Terminal
Test Panel
(Optional)
External Power Source
Virtual Backplane Bus (PCI)
CMC Module (running COTS Operating System)
Backplane Interface Controller Level A
Aircraft Fault History Central Maintenance Software Fault Processing - Level ED Aircraft Condition Monitoring - Level E I/O Processing - Level E
ACMF Data LDI Database
Time Limited Dispatch Aircraft Fault History
BIC
ACMF Data LDI Database
PROCESSOR MODULE BIC
(Optional)
Remote Image Bus (RIB) ARINC 429
DATABASE MODULE (Backup Data Files)
Data Management Unit
Copilot CCD
Copilot Multi-Function Display ASCB Member Systems
Time Limited Dispatch
NIC BIC
Ethernet LAN
ASCB-D
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Operation ACCESSING THE MAINTENANCE SYSTEM The remote terminal ADMS functionality and GUI is the same as that provided in the cockpit using the copilot’s MFD and CCD.In order to access the main CMCF menu in the cockpit, the following procedures should be followed: • Select the Maintenance menu for display on MFD No.2 by: - Using the CCD No.2 touch pad to move the cursor to the Maintenance Soft Key. • Select the Maintenance Soft Key by pushing one of the enter keys on CCD No.2: - The main Maintenance menu is displayed.
Figure 9: Accessing to CMC Maintenance pages
CCD – CURSOR CONTROL DEVICE
MFD 2
Maintenance System Config
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CMC Main Menu Through the CMC Main Menu, the following options can be accessed: • Maintenance Messages Display • System Diagnostics • Extended Maintenance • Data Loader
Figure 10: CMC Main Menu
MAINTENANCE M ESSA GES
SYSTEM DIA GNOSTICS
EXTENDED MAINTENA NCE
DATA LOADER
FILE TRANSFER
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Maintenance Messages Menu Select Maintenance Messages in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the Maintenance Messages Soft Key • Select the Maintenance Messages Soft Key by pushing one of the enter keys on CCD No.2 • The Maintenance Messages menu is displayed The active (with EICAS correlation if present),present leg, historical by date, and historical by ATA selections present a screen with a list of messages in the selected conditions.
• Historical by Date All stored maintenance messages whether they are active or inactive, shown sorted by date.Through this selection, the “SELECT A FLIGHT LEG” page is displayed and a list of dates is first presented.Once a date is selected a list of flight legs are presented.Selecting a flight leg, a list of FDE messages, organized by classification (warning, caution, advisory or status), are presented.Selecting a message, the “FDE DETAIL” menu is displayed showing the details of the related message. • Historical by ATA (Air Transport Association of America)
Maintenance Messages Display
All stored maintenance messages whether they are active or inactive are displayed and a list of FDE messages, organized by classification (warning, caution, advisory or status), are presented.Selecting a message, the “FDE DETAIL” menu is displayed showing the details of the related message.
The Maintenance Messages Display are classified as follow:
• Maintenance Message Details
• Active
Data in the “MAINTENANCE MESSAGES DETAIL” page, displayed from any of the above selections, including the fault name,type and code, a field for the LRUs at fault, a field for symptom text, a field for linked document if any, and a field for the maintenance message occurrences are shown in chronological order.
active maintenance messages show all fault messages.Selecting this option, the “ACTIVE CORRELATED FDE (Flight Deck Effect)” page is displayed and a list of FDE messages, organized by classification (warning, caution, advisory or status), are presented.Selecting a message, the “FDE DETAIL” menu is displayed showing the details of the related message. • Present Flight Leg All maintenance messages that occurred during the last flight leg flown.Selecting this option, the “STORED MAINTENANCE MESSAGES” page is displayed and a list of FDE messages, organized by classification (warning, caution, advisory or status), are presented.Selecting a message, the “FDE DETAIL” menu is displayed showing the details of the related message.
• CAS (Crew Alerting System) Message Correlation For every EICAS (Engine Indicating and Crew Alerting System) message with an associated maintenance action there is a corresponding maintenance message.For example, if the CMC fails, a corresponding CAS message will be set to EICAS:
Figure 11: CMC Maintenance Messages
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Maintenance Message Details • Maintenance Message Details Data in the “MAINTENANCE MESSAGE DETAIL” page, displayed from any of the above mentioned selections, including the fault name, type and code, a field for the LRUs at fault, a field for symptom text, a field for linked documents if any, and a field for the maintenance message occurrences are shown in chronological order.
Figure 12: CMC Maintenance Message Page
CMC Maintenance Message Page MAINTENANCE MES S AGE DETAIL F AULT N AME: WASTE WAS TE S SERV ERV PNL P NLSSW/WWS W/WWSC/MAU3 C/MAU3 FAULT FAULT F AULT TYP E : P ROBE / S ENS OR F AULT CODE: 38325784PNL
LRU (S ) AT F AULT: VWS S e rvice P a ne l S witch Wa te r & Waste Wa s teSSyste ys temmController Controlle r MAU3 – Ge ne ric I/O Module (s (slot lot 10) Aircra ft Wiring
LRUs/LRMs: highest to lowest probability
S YMP TO M: Che ck MAU3 for fault fa ult re reporting. porting. S e rvice pa ne l door door switch s witchmay m a ybe beimprope imprope y rly arldjus adjusted te d or m malfunctioning. a lfunctioning.
To add value for maintenance. Not mandatory
DOCUMENTS
Link to FIM (Fault Isolation Manual)
MAINTEN ANCE MES S AGE OCCURRENCES : ACTIVE ACTIVE INACTIVE IN ACTIVE ACTIVE ACTIVE
MAIN MENU
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System Diagnostics menu Select System Diagnostics in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the System Diagnostics Soft Key • Select the System Diagnostics Soft Key by pushing one of the enter keys on CCD No.2 • The System Diagnostics menu is displayed and a list of member systems organized by ATA chapter that have system diagnostic pages associated with them are presented.
Figure 13: CMC System Diagnostics Page
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Extended Maintenance The extended maintenance has the following options: • Storage enabled/disabled Inhibits storage of maintenance messages in the FHDB (Fault History Database). Selecting the “STORAGE ENABLED PRESS TO DISABLED” soft key, the soft key label will change to “STORAGE DISABLED PRESS TO ENABLE”.In order to re-enable storage press the soft key again. • Member System Status Supplies an indication, if the member system is operational or not. Through this option, a “MEMBER SYSTEM STATUS” screen is displayed listing the member systems organized by ATA chapter.The status for each can be shown as “OPERATIONAL”, “NOT OPERATIONAL”, “TEST”, “LRU NO COM” and “IOM (Input/Output Module) NO COM”. • Configuration Shows the equipment configuration of the member systems. Through this selection, a “SYSTEM CONFIGURATION” page is displayed.There are 30 fields for the display of related data that are pre-defined as:equipment Id (Identification), destination identifier, hardware part number, serial number, and software version.A list of member systems, organized by ATA chapter that transmit their system configuration, are presented. • Reports There are two options of reports: “CMC REPORTS” and “ACMF REPORTS”
The “CMC REPORTS” are classified in three types: “ACTIVE MAINTENANCE MESSAGE”, “CURRENT LEG FAILURES”, and “CONFIGURATION REPORT”.The RT or the DMU can be used to send these reports to storage or to the printer. Through the “CMC REPORTS” selection, a list of printable reports is displayed.Selecting one of three types above, a list of destinations where the reports can be sent is displayed (“LOCAL STORAGE”, *DMU 1 PCMCIA SLOT 1”, “DMU 1 PCMCIA SLOT 2” or “COCKPIT PRINTER”).The selected report will be sent and a report status field will be displayed in the “CMC REPORTS” page.The “EXPORT FAULT HISTORY” is also an option available in the CMC reports section to download the FHDB from the aircraft.The RT or the DMU can be used to send these reports only to storage.The “ACMF REPORTS” are trend and exceedence reports.Once stored, data can be retrieved using the RT or DMU or retrieved for in-air download via the CMF (Communications Management Function).
Figure 14: CMC System Diagnostics
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Member System Status menu Select Member System Status in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the Member System Status Soft Key • Select the Member System Status Soft Key by pushing one of the enter keys on CCD No.2 The Member System Status is displayed and a list of member systems organized by ATA chapter and the corresponding status is presented.
Figure 15: CMC Member System Status Page
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File Transfer Menu Select File Transfer in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the File Transfer Soft Key • Select the File Transfer Soft Key by pushing one of the enter keys on CCD No.2 The File Transfer menu is displayed and a list of member systems organized by ATA chapter that have file transfer associated with them are presented.
Figure 16: CMC File Transfer Page
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System Configuration Menu Select System Configuration in the Maintenance menu by: • Using the CCD No.2 touch pad to move the cursor to the System Configuration Soft Key • Select the System Configuration Soft Key by pushing one of the enter keys on CCD No.2 The System Configuration menu is displayed and a list of member systems organized by ATA chapter that transmit their system configuration are presented.
Figure 17: CMC Configuration Page
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Data Loader The Data Loader System (DLS) is a software program that allows the loading of the entire Primus Epic System. there are two loading modes. The “Target Load” allows the loading of a specific LRU and the “Full Load” allows the loading of the entire system. Selecting “FULL LOAD” or “TARGET LOAD” on the bezel, the “DLS INSTALLATION FUNCTION” are presented, showing the system drives. Selecting the drive and the file which will be the source of the load, the procedure is started.
Figure 18: Data Loader
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Load Times • • • • • •
Load times using Remote Terminal DLS: System Configuration and Status Check takes approx. 10 minutes or less Full System Load takes approx. 1:30 hrs. Full System Load including DB module 2:30 APM Load approx. 10 minutes or less Seperately Loadable Database: FMS NAV, A/C, COMPANY, CUSTOM: 30 minutes or less • Separately Loadable DAtabase: EGPWS Terrain and Envelope Modulation: approximately 30 minutes • LDI. 15 minutes
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Reports Menu • Reports - There are two options of reports.”CMC REPORTS” and “ACMF REPORTS”. The “CMC REPORTS” are classified in three types: “ACTIVE MAINTENANCE MESSAGES”, “CURRENT LEG FAILURES” and “CONFIGURATION REPORT”.The RT or the DMU can be used to send these reports to storage or to the printer. Through the “CMC REPORTS” selection, a list of printable reports is displayed.Selecting one of three types above, a list of destinations where the report can be sent is displayed (“LOCAL STORAGE”, “DMU 1 PCMCIA SLOT 1”, “DMU 1 PCMCIA 2” or “COCKPIT PRINTER”).The selected report will be sent and a report status field will be displayed in the “CMC REPORTS” page. The “EXPORT FAULT HISTORY” is also an option available in the CMC reports section to download the FHDB from the aircraft.The RT or the DMU can be used to send these reports only to storage. The “ACMF REPORTS” are trend and exceedence reports.Once stored, data can be retrieved using the RT or DMU or retrieved for in-air download via the CMF (Communications Management Function). DATA LOADER The Data Loader System (DLS) is a software program that allows the loading of the entire Primus Epic system.There are two loading modes.The “TARGET LOAD” allows the loading of a specific LRU and the “Full Load” allows the loading of the entire syste.Selecting “FULL LOAD” or “TARGET LOAD” on the bezel, the “DLS INSTALLATION FUNCTION” are presented, showing the system drives.Selecting the drive and the file which will be the source of the load, the procedure is started. IN-AIR DIAGNOSTIC DOWNLOADS During an in-air operation the CMCF supports in-air diagnostic downloads through ASCB to the CMF.Both CMCF and ACMF reports can be automatically transmitted.The requested down link communication path is hard-cod
ed to first available and no up link commands are supported by the CMCF or the ACMF. PRINTER The CMCF supports communications with the printer through the LAN in all phases of flight. • RT - Lets all CMCF operate from the RT, as well as access the on-line linked manuals. • On-line linked Maintenance Manuals - Accessed from the RT only.This function lets a hypertext link between a maintenance message and the related text within the supplied electronic manuals. File Transfer - Lets the file transfer from the member system to the DMU or local drives of the notebook PC running on the RT (e.g. member system NVM (Non-Volatile Memory) files).
Figure 19: CMC Reports Page
CMC REPORTS ACMF REPORTS CMC REPORTS ACTIVE FDE/MAINT MSG SEND TO -> COCKPIT PRINTER SEND TO -> DMU 1 PCMCIA SLOT 1 SEND TO -> DMU 1 PCMCIA SLOT 2 CURRENT LEG FDE/MAINT MSG SEND TO -> COCKPIT PRINTER SEND TO -> DMU 1 PCMCIA SLOT 1 SEND TO -> DMU 1 PCMCIA SLOT 2 SYSTEM CONFIGURATION SEND TO -> COCKPIT PRINTER SEND TO -> DMU 1 PCMCIA SLOT 1 SEND TO -> DMU 1 PCMCIA SLOT 2 EXPORT FAULT HISTORY SEND TO -> DMU PCMCIA SLOT 1 SEND TO -> DMU PCMCIA SLOT 2 REPORT STATUS:
MAIN MENU
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Fault History Database All fault occurrences are stored in the Fault History Database (FHDB) which is located in the CMC module and implemented as three data files to prevent corruption.The FHDB is limited to 10 MB total which corresponds to an expected average year’s worth of aircraft fault storage.A maintenance message record is equivalent to 57 bytes.Once the storage limit is reached the oldest records are deleted first. FHDB redundancy and aircraft fault history are preserved by maintaining a copy of the FHDB in the aircraft Database module.Therefore if a CMC module is replaced the fault history of the aircraft is preserved.
Figure 20: CMC - Screen Hierarchy
NO CMC INTERF
In-- Flight In
ACTIVE MAINT MSG
On--Ground On
EXTEND MAINT
MAINT MSG MENU
PRES LEG MSG
CMC MAIN MENU
SYSTEM DIAG
HIST MAINT MSG BY DATE
HIST MAINT MSG BY ATA
STORAGE ALLOWED
MS STATUS
SYSTEM CONFIG
REPORTS
TEST PAGES
TO AIRCRAFT
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MS STATUS
TO STORAGE
TO DMU
TO PRINTER
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45-MEL (Example) ------------------------------------------------------------------------------¦ U.S. DEPARTMENT OF TRANSPORTATION ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
MASTER MINIMUM EQUIPMENT LIST FEDERAL AVIATION ADMINISTRATION --------------------------------------------------------------------------AIRCRAFT: ¦ REVISION NO: 2 ¦ PAGE: ERJ-170, ERJ-190 ¦ ¦ ¦ DATE: 11/16/2004 ¦ 45-1 --------------------------------------------------------------------------1. ¦ 2. NUMBER INSTALLED SYSTEM & ¦ -------------------------------------------SEQUENCE ITEM ¦ ¦ 3. NUMBER REQUIRED FOR DISPATCH NUMBERS ¦ ¦ --------------------------------------------------------------- ¦ ¦ ¦ 4. REMARKS OR EXCEPTIONS 45 CENTRAL MAINTENANCE ¦ ¦ ¦ COMPUTER ¦ ¦ ¦
¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦ ¦
¦ 45-01 Central ¦ Maintenance ¦ Computer (CMC)
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ May be inoperative provided ¦ procedures do not require its use. ¦
¦ ¦ ¦
¦ 45-03 Data-Loader ¦ *** Management Unit ¦ (DMU)
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ (M)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
¦ ¦
D ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
¦ ¦
¦ 45-04 Database (DB) ¦ Module ¦
C ¦ 1 ¦ ¦
¦ 0 ¦ ¦
¦ (M)May be inoperative provided ¦ alternate procedures are ¦ established and used.
¦ ¦ ¦
¦ ¦
D ¦ 1 ¦
¦ 0 ¦
¦ May be inoperative provided ¦ procedures do not require its use.
