Continental Airlines
B737-300/-500
Electrical and Avionics
NOTICE
THIS MANUAL HAS BEEN PREPARED FOR AIRPLANE SYSTEMS AND
AVIONICS TRAINING. IT WILL NOT BE REVISED AND DOES NOT AMEND
OR SUPERSEDE INFORMATION CONTAINED IN APPLICABLE
GOVERNMENT REGULATIONS AND CONTINENTAL'S APPLICABLE
SERVICE BULLETINS, MAINTENANCE MANUALS, OVERHAUL
MANUALS AND WRITTEN INSTRUCTIONS.
CONTINENTAL AIRLINES MAINTENANCE TRAINING DEPARTMENT
A
TABLE OF CONTENTS
VOLUME 1
SUBJECT
PAGE
GENERAL INTRODUCTION
1 - 60
ELECTRICAL POWER
1 - 90
ARINC 429 AND ABBREVIATIONS
1 - 28
EFIS
1 - 100
AIR DATA SYSTEM
1 - 74
INERTIAL REFERENCE SYSTEM
1 - 122
AUTOMATIC DIRECTION FINDER
1 - 24
VHF NAVIGATION
1 - 64
MARKER BEACON
1 - 10
LOW RANGE RADIO ALTIMETER
1 - 32
DISTANCE MEASURING EQUIPMENT
1 - 30
GROUND PROXIMITY WARNING SYSTEM
1 - 72
INSTRUMENT COMPARATOR
1 - 14
VOLUME 2
ATC TRANSPONDER
1 - 30
TCAS
1 - 54
WEATHER RADAR
1 - 66
B
SUBJECT
PAGE
FLIGHT MANAGEMENT SYSTEM INTRODUCTION
1 - 10
FLIGHT MANAGEMENT COMPUTER SYSTEM
1 - 92
DFCS INTRODUCTION
1 - 28
DFCS HYDRAULICS AND FLIGHT CONTROLS
1 - 30
DFCS OPERATIONS
1 - 62
DFCS ENGAGE AND WARNING MONITORS
1 - 18
DFCS PITCH AND ROLL CONTROL
1 - 42
DFCS BITE
1 - 18
DFCS EFIS DIFFERENCES
1 - 30
YAW DAMPER
1 - 22
AUTOTHROTTLE
1 - 50
CATEGORY II/AUTOLAND
1 - 16
AUDIO INTEGRATING
1 - 30
SERVICE INTERPHONE SYSTEM
1 - 16
PASSENGER ADDRESS SYSTEM
1 - 30
VHF COMMUNICATIONS
1 - 20
ACARS
1 - 46
CREW CALL SYSTEM
1-8
COCKPIT VOICE RECORDER
1 - 14
DIGITAL FLIGHT DATA RECORDER
1 - 60
CLOCKS
1 - 12
FUEL QUANTITY
1 - 22
C
General
Introduction
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GENERAL DESCRIPTION
INSTRUMENT PANEL LOCATIONS
1. General
Flight instruments and avionics system controls and indicators are located on panels in the
flight compartment.
2. Components
Flight instruments and annunciator lights are located on the captain's, first officer's, and
center instrument panels. Control panels and annunciator lights are located on the FWD and
AFT overhead panel, light shield and sidewalls.
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INSTRUMENT PANEL LOCATION
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GENERAL DESCRIPTION
PANEL INSTRUMENT INSTALLATION
1. General
Most of the airplane's instruments are secured to panels by clamps. The clamps are fastened
to the back of the panel face and surround the instrument case. Loosening a clamp
facilitates instrument removal; tightening a clamp holds the instrument firmly in its panel
mounted position.
2. Clamp Mounting Screws
The clamps are fastened to the panels by clamp mounting screws which can be identified by
noting that they are smaller than the other screws positioned around the premimeter of the
instrument. Limited counterclockwise rotation of the clamp mounting screws facilitates
instrument removal after the clamp has been slackened. The clamp mounting screws should be
removed only when it is necessary to detach the clamp from the panel. On retangular
instrument installations, there are two clamp mounting screws located diagonally opposite
each other. On cylindrical instrument installations, there is one clamp mounting screw.
3. Clamp Adjustment Screws
A clamp may be loosened or tightened by turning the clamp adjustment screws which can be
identified by noticing that they are larger than the clamp mounting screws.
Counterclockwise rotation of the adjustment screws extends the clamp to permit removal or
re-orientation of the instrument; clockwise rotation contracts the clamp around the
instrument case. On retangular instrument installations, there are two clamp adjustment
screws located diagonally opposite each other. On cylindrical instrument installations,
there is one clamp adjustment screw.
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RECTANGULAR INSTRUMENT INSTALLATION
ROUND INSTRUMENT INSTALLATION
PANEL INSTRUMENT INSTALLATION
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NOTES
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CAPTAINS INSTRUMENT PANEL
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PILOTS CENTER INSTRUMENT PANEL AND LIGHTSHIELD
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Center Panel
EFFECTIVITY
CAL 320-324, 336-345
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FIRST OFFICERS INSTRUMENT PANEL
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PILOTS' CONTROL STAND - DESCRIPTION AND OPERATION
Pilots' Control Stand
EFFECTIVITY
CAL 301-361
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PILOTS ELECTRONICS PANEL
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FORWARD OVERHEAD PANEL - DESCRIPTION AND OPERATION
Forward Overhead Panel
EFFECTIVITY
CAL 301-319, 325-335, 346-361
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Forward Overhead Panel
EFFECTIVITY
CAL 320-324, 336-345
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AFT OVERHEAD PANEL - DESCRIPTION AND OPERATION
Aft Overhead Panel
EFFECTIVITY
CAL 301-361
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AUXILIARY PANELS - DESCRIPTION AND OPERATION
Auxiliary Panels
EFFECTIVITY
CAL 301-330
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Auxiliary Panels
EFFECTIVITY
CAL 331-361
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FORWARD CABIN ATTENDANTS' PANELS - DESCRIPTION AND OPERATION
Forward Cabin Attendants' Panels
EFFECTIVITY
CAL 301-361
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AFT CABIN ATTENDANTS' PANELS - DESCRIPTION AND OPERATION
Aft Cabin Attendants' Panels
EFFECTIVITY
CAL 301-361
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MAIN PANEL-CAPTAIN
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Center Panel
EFFECTIVITY
CAL 380-386, 601-699
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Pilots' Lightshield
EFFECTIVITY
CAL 638-699
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MAIN PANEL-FIRST OFFICER
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Pilots' Control Stand
EFFECTIVITY
CAL 380-386, 601-699
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Aft Electronic Panel
EFFECTIVITY
CAL 601-633, 638-699
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Forward Electronic Panel
EFFECTIVITY
CAL 601-633
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Forward Electronic Panel
EFFECTIVITY
CAL 380-386, 634-699
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Forward Overhead Panel
EFFECTIVITY
CAL 601-618
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Forward Overhead Panel
EFFECTIVITY
CAL 619-633, 637-699
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Aft Overhead Panel
EFFECTIVITY
CAL 380-386, 601-699
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Forward Cabin Attendants' Panels
EFFECTIVITY
CAL 380-386, 601-699
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Aft Cabin Attendants' Panels
EFFECTIVITY
CAL 380-386, 601-699
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LOAD CONTROL CENTER
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GENERAL DESCRIPTION
MASTER WARNING LIGHTS AND CAUTION ANNUNCIATORS
1. General
The master caution lights inform the pilots that a system fault indicator light has
illuminated on overhead panel or the engine and APU protection panel. The system caution
annunciator informs the pilots which system has a fault.
2. Location
The master caution lights and system annunciators are located on the light shield.
3. Power
Power for the lights is either from a system circuit breaker or from the DIM and TEST
circuit breaker.
4. Operation
The master caution light consists of an amber light cap labeled MASTER CAUTION - PUSH TO
RESET.
Each system caution annunciator is composed of an amber light cap divided into six sections.
The light cap sections on the captain's side are labeled FLT CONT, ELEC, IRS, APU, FUEL and
OVHT/DET. The light cap sections on the first officer's side are labeled ANTI-ICE, ENG,
HYD, OVERHEAD, DOORS, and AIR COND.
Pressing either the captain's or first officer's master caution light extinguishes the
master caution lights and the system caution annunciators. The system fault indicator light
remains illuminated until the fault is corrected. The pilot may recall system fault
indications or the system caution annunciators by pressing either annunciator. All
annunciator lights will illuminate while the annunciator is pressed. When released, only
those system with illuminated fault indicators will cause the corresponding system caution
annunciator to remain illuminated.
The lamps for the warning lights and annunciators are replaced by pulling off the light cap.
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MASTER WARNING LIGHTS AND CAUTION ANNUNCIATORS
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COMPONENT FUNCTIONAL DESCRIPTION
ANNUNCIATOR IDENTIFICATION
The two annunciators are composed of six sections, each with individual lights. Several systems
are circuited to each section. A section illuminates when one or more of the systems has a
failure.
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ANNUNCIATOR IDENTIFICATION
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ELECTRONIC EQUIPMENT COMPARTMENT
ELECTRONIC EQUIPMENT RACK LOCATIONS
1. General
The majority of avionic system line replaceable units (LRUs) are located on the equipment
racks in the electronic equipment (E&E) compartment.
2. Location
The E&E compartment is located in the lower 41 section, aft of the nose wheel well.
3. Components
There are 3 equipment racks located in the E&E compartment. The external access door is
located between the 3 racks. Access from the forward cargo compartment is through a
removable panel located between the E2 and E3 racks. There is an additional access to the E&E
compartment from the passenger cabin. This access is located between the seat tracks on the
right side of the airplane by the third window from the front. This is not a normal access
since there are two panels, one must be removed from below and one from above.
ELECTRONIC EQUIPMENT RACK LOCATIONS
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ELECTRONIC EQUIPMENT RACKS
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ELECTRONIC EQUIPMENT COMPARTMENT
ELECTRONIC EQUIPMENT RACK LOCATIONS
General
Most avionic systems line replaceable units (LRUs) are located on the equipment racks
in the electronic equipment (E/E) compartment.
Location
The E/E compartment is located in the lower 41 section, aft of the nose wheel well.
Components
There are 4 equipment racks located in the E/E compartment. The external access door
is located in the floor between the 4 racks.
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ELECTRONIC EQUIPMENT RACK
MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
MODULE INSTALLATION
General Features
Module retaining devices provide the means of pushing the module connector into its mating
rack connector, securing on the shelf, and extracting the unit for removal. There are three
types of module hold downs: two are thumbscrew types, the other is a lever-latch.
Cam-Lock Lever
The cam-lock lever is a force limiting, latching lever which also serves as a carrying
handle for the LRU after removal. Press the latch to release the lever from the
handle. Move the lever downward to release the module from the connector. To install
the module, slide the unit into the shelf with the lever open until the lever
engages the shelf-mounted fork. Then move the lever upward to its latched position.
Thumbscrew
Thumbscrew hold downs are retainer cups which are placed over module hold down hooks and
tightened using the retainer nuts. Because this type of hold down does not provide the
necessary force limiting, an extractor tool must be used for removal and installation.
The extractor tool is placed under the front edge of the module. The torque knob is
rotated to release the unit from the connector. Rotating the knob in the opposite
direction will apply the necessary force to insert the tool into the shelf. The tool
is stowed in a bracket on the forward, inboard stanchion of the E2 rack.
Tridar Extractors
The tridar extractors are torque limiting, thumbscrew retaining devices which include
an extracting function. They also provide a visual indication when the hold down is
applying the desired force. To remove a module, loosen the extractor knob until the
red band is visible. Rotate the keeper to align the slot with the T-hook and free the
extractor. To install the module, align the slot with the T-hook and rotate the keeper
to the wide band. Tighten the knob until the red band is out of view.
NOTE:
Some modules are static sensitive and contain devices that can be damaged by
static discharge. Do not handle modules labeled static sensitive before
reading procedures for handling electrostatic discharge sensitive devices.
(Reference Maintenance Manual, Chapter 20-41-01.)
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MODULE INSTALLATION
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GENERAL DESCRIPTION
ELECTROSTATIC DISCHARGE SENSITIVITY
General Subsystem Features
Static Build-Up
Static electricity is generated and stored on the surface of
non-conductive materials and discharges to the first available ground source. Items
such as human hands, air, and glass store high positive charges, whereas plastics store
large charges of negative electricity.
Static Damage
If static discharge can be seen or felt, then it may be assumed that the potential
difference prior to discharge can be measured in thousands of volts. However, the
voltage necessary to damage microcircuitry can be thirty volts or less. Therefore,
electrostatic discharge damage can occur even though the discharge is of insufficient
strength to be felt or seen. The low energy source that most commonly destroys ESDS
devices is the human body which, in conjunction with nonconductive garments and floor
coverings, generates and retains static electricity. In order to adequately protect
ESDS devices, the device and everything that contacts it must be brought to ground
potential by providing a conductive surface and discharge paths.
Handling
LRU's/INSTRUMENTS – Do not touch connector pins or other exposed conductors. Install
dust caps whenever a unit is removed.
Circuit Cards – Use a wrist strap assembly when handling circuit cards. Wrap card in a
conductive plastic bag and place in a container for transport.
Labeling
Racks
Racks and shelves containing ESD sensitive equipment are identified.
Equipment/Instruments
LRU's and instruments containing ESD sensitive circuitry are labeled.
Circuit Cards
Circuit assemblies containing ESD sensitive equipment are labeled accordingly on
the extractor lever.
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STATIC DISCHARGE SENSITIVE DEVICES IDENTIFIERS
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ELECTRONIC EQUIPMENT COMPARTMENT
EQUIPMENT COOLING
1. General
The equipment cooling system cools the equipment on the racks in the electronic compartment,
some circuit breaker panels in the flight compartment and the main instrument panel. Cabin
air is the cooling medium. It is drawn through and around the equipment into a system of
ducts and manifolds.
2. Location
The system components are located in the lower aft section of the electronic equipment
compartment. The blower select switch and cooling OFF warning light are on the overhead
panel (P5).
3. Components
A.
Blowers
The main blower operates continuously when power is supplied to the airplane. If the
main blower fails, the alternate blower may be switched on. Check valves prevent
recirculating air being discharged from the operating blower back through the nonoperating blower.
B.
Automatic Flow Control Valve
This valve controls the overboard flow of air as the cabin-to-ambient differential
pressure increases.
C.
Equipment Mounting Shelves and Ducts
The electronic racks are formed by shelves, header plenums, and supporting structure.
Holes for airflow are suitably located for each shelf.
D.
Airflow Detection System
The airflow detection system provides a warning when there is insufficient cooling air.
The airflow detection system consists of an equipment cooling off light and system
circuitry.
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EQUIPMENT COOLING COMPONENT LOCATION
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NOTES
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ELECTRONIC EQUIPMENT COMPARTMENT
EQUIPMENT COOLING OPERATION
1. General
Two blowers, main and alternate, are provided for the equipment cooling. The operating
blower is selected by a switch on the P5 overhead panel. The blower operates, when power is
applied to the electrical buses.
2. Operation
The operating blower draws cabin air through and around the equipment and into the plenums
and ducts. An airflow detector senses the operation of the blower, and if there is no
airflow, the Equipment Cooling OFF light on the overhead panel is illuminated.
When the airplane is on the ground, or during low altitude flight, the heated air is
discharged overboard. As differential pressure increases, the airfoil shaped Automatic Flow
Control Valve closes, directing the air to the forward cargo compartment. The valve closes
at 2.0 to 2.8 psi differential pressure.
EQUIPMENT COOLING OPERATION
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ELECTRONIC EQUIPMENT COMPARTMENT
EQUIPMENT COOLING SCHEMATIC
1. General
Equipment cooling operates when ac is applied to the airplane's buses.
2. Power
Equipment cooling requires 115 volts ac for the blower, and 28 volts dc for the warning
circuit. The circuit breakers are located on the P18 panel.
3. Operation
With power provided to the buses, a relay energizes, applying power to the selected blower.
28 volts dc is applied to the Airflow Detector. With normal airflow, the heater is cooled.
When airflow is low, temperature rises closing the switch. Equipment Cooling OFF, Master
Caution, and the OVHD annunciator illuminate. On the ground, the Crew Call Horn will also
sound, if either inertial reference unit is on.
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EQUIPMENT COOLING SCHEMATIC
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ELECTRONIC EQUIPMENT COMPARTMENT
EQUIPMENT COOLING
General
Equipment cooling is provided by separate supply and exhaust systems using cabin air as the
cooling medium. Supply fan cooling (blow through) is used for the Electronic Flight
Instrument System (EFIS) display units in the flight compartment and the inertial reference
units (IRU) in the electronic equipment E/E compartment. Exhaust fan cooling (draw through)
is used for the other equipment on the E/E racks, some circuit breakers in the flight
compartment and the main instrument panel.
Location
The components for the supply and exhaust systems are located in the E/E compartment except
for most of the ducting and the low flow detectors. Selector switches for normal and
alternate supply and exhaust fans and OFF warning lights are on the overhead panel (P5).
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EQUIPMENT COOLING
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ELECTRONIC EQUIPMENT COMPARTMENT
EQUIPMENT COOLING OPERATION
General
Normal and alternate supply fans and normal and alternate exhaust fans are provided for
equipment cooling. The operating supply and exhaust fans are selected by switches on the P5
overhead panel. The fans operate when power is applied to the electrical buses.
Operation
Supply Fan Cooling
The operating supply fan draws ambient air from the electronic equipment compartment
through the air cleaner and supplies air to the Inertial Reference Units (IRU) and the
main duct. Air passing through the main duct flows across the low flow detector probe
and into the four EFIS displays. The supply fan forces cooling air through the four
EFIS displays and out the exhaust ducts. This exhaust air is sent to the lower lobe of
the airplane. If the low flow detector senses a loss of cooling, the supply OFF light
on the P5 overhead panel and the master caution light illuminates.
Exhaust Fan Cooling
The operating exhaust fan draws cabin air through and around the equipment and into the
plenums and ducts. A low flow detector monitors the operation of the blower, and if
there is low airflow, the Equipment Cooling Exhaust OFF light on the overhead panel
illuminates.
When the airplane is on the ground, or at low altitude, the heated air is discharged
overboard. As differential pressure increases, the airfoil shaped Automatic Flow
Control Valve closes, directing the air to the forward cargo compartment. The valve
closes at 2.0 to 2.8 psi differential pressure.
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EQUIPMENT COOLING OPERATION
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ELECTRONIC EQUIPMENT COMPARTMENT
SUPPLY FAN COOLING SCHEMATIC
General
Supply fan cooling operates when 115v ac power is applied to the airplane's buses
Power
Supply fan cooling requires 115v ac for the fans and 28v dc for relay and low flow detector
operation. Circuit breakers for 115v ac and 28v dc for the relays and low flow detector are
located on P18. The circuit breaker for the 28v dc to the supply fan OFF light is on the
P6.
Operation
When 115v ac power is applied to the airplane's buses, a relay in the P6 panel energizes,
applying power to the selected supply fan. The low flow detector is a temperature sensing
probe assembly designed to monitor the cooling air temperature. The low flow detector
provides an alarm signal when the cooling air temperature rises above a temperature sensing
threshold. It is powered by 28v dc. The airflow through the equipment cooling system
removes heat from the detector. When the threshold is exceeded, a low flow signal is sent
to the EFIS symbol generators and illuminates the OFF, MASTER CAUTION and OVERHEAD lights.
The OFF light remains on until airflow through the system is sufficient to remove heat from
the detector. The crew call horn will sound if the airplane is on the ground and either IRU
is powered on.
The normal and alternate supply fan motors are protected by heat-sensitive switches. There
is one thermal switch per motor winding. When the motor temperature reaches 400°F, the
switch opens. This de-energizes the controlling relay, removing power to the winding.
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SUPPLY FAN COOLING SCHEMATIC
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STATIC DISCHARGERS
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GENERAL DESCRIPTION
ANTENNA LOCATIONS
General Component Locations
Antennas associated with the communications and navigation systems
are shown on the graphic below.
The weather radar, VHF NAV glide slope, and dual localizer antennas
are mounted in the radome. There are VHF COMM 1, TCAS, and ATC-1
antennas located on top of the fuselage.
The VHF navigation
VOR/LOC antenna is incorporated into the tip of the vertical fin.
All other system antennas are situated on the bottom of the
fuselage.
ANTENNA LOCATIONS
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NOTES
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Electrical
Power
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NOTES
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INTRODUCTION
ELECTRICAL POWER DIAGRAM
1. Purpose
The electrical power is used for the control, operation, and indication of the various
airplane systems on the ground and inflight.
2. System Description
The power is obtained from the battery, generation and ground support equipment. It is
controlled and monitored prior to distribution as 115 volt ac, 28 volt ac and 28 volt dc supply
to using systems.
ELECTRICAL POWER
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GENERAL DESCRIPTION
ELECTRICAL POWER
General Component Locations
The alternating current power is supplied by two engine driven generators for normal
inflight operation. A generator, driven by the auxiliary power unit, can supply all power
for ground or flight operation. Power can be supplied on the ground through the external
power (ac) receptacle. The dc power is supplied from the battery or conversion from the ac
power.
The ac and dc power is distributed to the various systems through the left and right load
control centers in the flight compartment.
The control and indication of the electrical system is from the P5 overhead panel in the
flight compartment.
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ELECTRICAL POWER
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INTRODUCTION
GENERATOR DRIVE
1. Purpose
The ac power produced by the generators is of constant frequency. To achieve this the
generators are rotated at constant speed by the drives on the engines and the APU.
2. System Description
The drive system consists of constant speed drives (CSD), an auxiliary power unit with speed
reduction gearbox, controls and monitoring system.
GENERAL DESCRIPTION
1. General Component Locations
The constant speed drives are located on the front left side of each engine, the APU is in
the aft section of the fuselage, and the control and monitoring is from the P5 panel in the
flight compartment.
2. General Operation
The engine generator drives are hydromechanical units increasing or decreasing the engine
input speed and supplying constant speed to the generators. Oil is used as an operating
fluid. The APU operates at constant speed and through the gearbox the speed is reduced for
the generator.
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ENGINE GENERATOR DRIVE
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COMPONENT FUNCTIONAL DESCRIPTION
CONSTANT SPEED DRIVE
Physical Description
The CSD is a hydromechanical transmission using hydraulic trimming and mechanical controls
to govern the output rotational speed.
Located on the CSD are:
Oil inlet and outlet connection with temperature bulbs
Wet spline cavity fill valves
Main electrical connector and harness.
Vent valve mount
Low oil pressure switch
Oil quantity sight glass
Disconnect solenoid
Output pad drain (Generator Side)
Charge filter with differential pressure indicator
Drain plug (Reservoir)
Case drain and magnetic drain plug
Governor adjustment
Reset handle
On the input end of the CSD is a QAD (Quick Attach Detach) flange for CSD attachment to the QAD
ring on the gearbox. On the opposite face are 12 studs for the generator attachment.
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CONSTANT SPEED DRIVE
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COMPONENT FUNCTIONAL DESCRIPTION
CONSTANT SPEED DRIVE SCHEMATIC
Operation
Hydraulic system
The constant speed drive consists of two positive displacement axial slipper pistontype hydraulic units and a mechanical axial geared differential which performs the
speed summing function. One hydraulic unit is of fixed displacement and the other is a
variable displacement type.
The hydraulic system consists of the charge pump, the scavenge pump and the charge relief
valve. The charge pump is located in the system between the all-attitude reservoir and the
transmission. It supplies oil to the hydraulic units, governor, control piston and the
lubricating system.
The charge pump forces the oil through the charge filter which incorporates a pressure
differential indicator. The red button in the indicator will pop out at 45-55 psi
differential pressure. Connected to the charge line is a low oil pressure switch which
operates at 120-160 psi.
The scavenge pump is located in the system between the transmission sump and the external
oil cooler. The scavenge pump picks up lube oil and internal leakage and pumps it through
the external oil cooler into the all-attitude reservoir.
Temperature Control
Heat produced by the transmission is absorbed by the hydraulic fluid in the unit and
dissipated in an oil cooler at the forward end of the engine. The hot oil passes
through an external filter then through the cooler. Oil temperature is sensed at the
outlet and inlet from the CSD. Normal temperature rise of the oil through the
transmission is about 10°C at continuous full load, with an inlet oil temperature of
about 120°C, at normal input speeds. The oil in the transmission serves as a
lubricant, as a coolant and as the hydraulic medium of the drive.
The oil returning from the cooler is deaerated in a swirl chamber in the reservoir. If the
temperature of the oil reaches 157°C, a thermoswitch in the return line closes
illuminating the HIGH OIL TEMP light on the P5 panel.
A vent valve is provided to prevent negative pressure and retain positive pressure. Oil
quantity is indicated by a sight gage on the left side.
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CONSTANT SPEED DRIVE SCHEMATIC
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COMPONENT FUNCTIONAL DESCRIPTION
CONSTANT SPEED DRIVE-ELECTRICAL
Monitor
The constant speed drive has electrical components connected to external circuits for
monitoring in the flight compartment on the constant speed drive and standby power panel P55. The power supplies are from the circuit breaker panel, P6.
The low oil pressure switch operates at 120-160 psig. With increasing pressure, the switch
opens at 160 psig; and with decreasing pressure, the switch closes at 120 psig. The switch
controls the LOW OIL PRESSURE amber light.
At a temperature of 157°C, the high oil temperature switch closes and the HIGH OIL TEMP
amber light illuminates.
The disconnect solenoid is activated by 28 volt dc from a guarded switch on the P5 panel.
The normally closed contacts of the switch connect a ground to the disconnect solenoid; thus
both sides of the coil are grounded to prevent the possibility of voltage pickup and
inadvertantly tripping the CSD. A second section of this switch provides 28 volt dc to the
trip coil of the generator control relay, tripping off the generator field.
Two variable resistance temperature bulbs measure the oil temperature on either side of the
oil cooler. One bulb measures the input oil temperature to the CSD and is read on a meter
on the P5 panel. A switch on the panel alters the circuit to include the oil out
temperature bulb so that the meter can indicate the rise in oil temperature through the CSD.
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MAINTENANCE TRAINING MANUAL
CONSTANT SPEED DRIVE ELECTRICAL
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
CSD OIL COOLER
1. Purpose
The CSD oil cooler is used for cooling of the oil from the CSD.
2. Location
It is located in an opening on the right side of the engine fan duct.
3. Physical Description.
The cooler is a small oil/air heat exchanger using fan air. A pressure relief valve is
installed in the cooler assembly.
4. Operation
During engine operation, hot oil from the CSD passes through the cooler; fan air provides
the cooling for the oil. If the cooler becomes blocked on the oil side, a pressure relief
valve cracks open at 50 psi ∆P and is fully open at 90 psi DP allowing oil to bypass the
cooler core.
5. Maintenance Practices.
The cooler can be removed by disconnecting the two oil lines and removing 8 nuts.
CAUTION:
COOLER DAMAGE CAN OCCUR FROM FOREIGN OBJECTS BEING INGESTED BY
THE ENGINE.
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MAINTENANCE TRAINING MANUAL
CSD OIL COOLER
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MAINTENANCE TRAINING MANUAL
CONSTANT SPEED DRIVE
Controls and Indicators
An amber LOW OIL PRESSURE light controlled by an electromagnetic pressure sensor
located on the CSD. The MASTER CAUTION and ELEC annunciator lights on the P7 panel
also illuminate.
An amber HIGH OIL TEMP light controlled by a thermoswitch located in the oil cooler
return line. The MASTER CAUTION and ELEC annunciator lights on the P7 panel also
illuminate.
A double-scaled generator drive oil temperature indicator. The indicator is wired into a
bridge circuit with two resistance type temperature bulbs. One scale displays temperature IN to
the CSD and the other scale displays RISE. The reading selection is controlled by a switch
located in the center on the panel. Both indicators are calibrated in degrees Celcius.
Two red, guarded, safety-wired, disconnect switches. Each switch controls a disconnect
solenoid in the respective CSD. Activation of the a switch disconnects the CSD mechanical
drive from the engine accessory gearbox drive.
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MAINTENANCE TRAINING MANUAL
Drive and Generator Attachment
The constant speed drive and the ac generator can be removed and installed as a
complete assembly, or as separate units. If these operations are being performed
without the aid of a cradle assembly, it will be easier to remove and install
them separately. A quick-attach-detach (QAD) locking ring facilitates removal
and installation of the CSD, or of the CSD and generator assembly.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
CONSTANT SPEED DRIVE FILTERS
Maintenance Practices
Filters
The constant speed drive oil system has two oil filters: an external filter on the
right side of the engine, and a CSD internal filter. Both filters have a differential
pressure pop-out indicators. When the filters are blocked, at a differential pressure
of 45 to 55 psid, the indicators pop-out.
General rules when changing filters:
1. Remove and discard old O-rings and install new O-rings.
2. Assure filter housings and screws are lockwired correctly.
3. Service the CSD to correct levels.
4. Motor the engine and check for leaks.
Warning/Caution
A. WARNING:
USE EXTREME CARE WHEN DRAINING CSD OIL OR REMOVING CSD COMPONENTS. HOT OIL
CAN CAUSE INJURY.
PROLONGED CONTACT WITH CSD OIL CAN CAUSE DERMATITIS. OIL WILL STAIN
CLOTHING AND CAN SOFTEN PAINT.
B. CAUTION:
TAKE ALL POSSIBLE PRECAUTIONS TO PREVENT FOREIGN MATTER FROM ENTERING DRIVE
SYSTEM.
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CONSTANT SPEED DRIVE FILTERS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
CSD AND WET SPLINE CAVITY SERVICE
1. Maintenance Practices
A.
Oil Servicing
The need and time for oil changes depends on operating conditions. An oil sight gage
on the CSD indicates the oil level in the system. A standpipe attached to the engine
accessory gearbox limits the oil level in the wet spline cavity.
B.
CSD Servicing
Service Constant Speed Drive Case and Reservoir
Step 1
If oil level is below bottom edge of band, service to bottom edge of band.
Pump oil (with 5 to 30 psi pressure) from service cart through CSD oil
reservoir pressure fill valve. Oil quantity is 4 1/2 quarts.
Step 2
Check oil level at reservoir oil sight gage. The operating range is within
the applicable band.
NOTE:
Oil level reading on sight gage of a disconnected drive may indicate
more oil than there actually is due to expanded air that may have
forced oil from oil cooler into transmission. To obtain a true oil
level reading on sight gage, pressure fill transmission, then
recheck sight gage.
Left and right bands are provided to allow for difference in angular
rotation of drive on No. 1 or 2 engine.
C.
Wet Spline Cavity Servicing
Step 1
Check oil level in the wet spline cavity by removing overflow drain cap on
engine gearbox and allow oil to flow from standpipe indicator inside the
cavity. When oil stops flowing spline cavity is filled to proper level. If
excessive oil flows, check either CSD or engine seals for oil leaking into
the wet spline cavity.
Step 2
To bring oil to overflow level, pump oil from service cart through spline
cavity fill check valve on CSD until it begins to flow from the standpipe
indicator. Oil quantity is 1 1/2 quarts.
Step 3
Allow excess oil to flow from standpipe, then replace overflow drain cap on
engine gear case.
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CSD WET SPLINE CAVITY SERVICE
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
CSD SPEED ADJUSTMENT
Maintenance Practices
Governor
The governor in the constant speed drive ensures that the output speed to the generator
is maintained at 6000 RPM corresponding to electrical frequency of 400 Hertz. If the
frequency is outside of this range, the governor can be adjusted.
Adjustment
Step 1
Obtain accurate frequency reading from the M400 test module. (The engine
should be operated for three minutes to allow the CSD to reach operating
temperature.) Stop engine and gain access to CSD.
Step 2
Remove adjustment cap to gain access to adjustment screw.
Step 3
Insert screwdriver in adjustment screw slot, press inward and turn
screwdriver until adjustment screw engages governor adjustment screw.
NOTE:
When adjustment screw is engaged with the driving governor
adjustment screw, it will be possible to feel a stop spring
ratcheting against a gear on the governor adjustment screw.
Step 4
Turn adjustment screw clockwise to raise frequency, counterclockwise to
lower frequency. One full turn equals approximately 14 Hertz.
Step 5
Adjust as necessary to bring frequency to 400 (±4) Hertz.
Step 6
Replace adjustment cap and O-ring over adjustment screw.
NOTE:
If filter or magnetic plug show indication of failure or
contamination, or if frequency is unstable or erratic, or if
frequency goes out of limits and stays out upon application or
removal of loads, replace CSD rather than adjust governor. If
filter, magnetic plug, or CSD is contaminated, also replace oil
cooler and associated tubing
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CSD SPEED ADJUSTMENT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
AC GENERATOR - ENGINE
1. Purpose
The generator driven by the engine produces the ac power required by the airplane systems
primiarly for flight operation.
2. Location
The generator is located on the forward left side of the engine and is attached to the
constant speed drive.
3. Physical Description
There are 6 ac and 2 ac terminals mounted on the generator.
T1, T2, and T3 are the power terminals; T4, T5 and T6 are the neutral or grounded
terminals.
T1 and T4 are for phase A, the leads are color-coded red.
T2 and T5 are for phase C, the leads are color-coded blue.
T3 and T6 are for phase B, the leads are color-coded yellow.
F and A are supply and return dc power for the generator field or excitation.
Cooling air enters through a duct from the engine fan duct, is collected in a shroud on the
generator and discharges overboard through a hole in the engine cowling (left side).
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AC GENERATOR - ENGINE
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
AC GENERATOR - APU
1. Purpose
The generator driven by the APU produces the ac power required by the airplane systems
primarily for ground operation. The generator can also be used as an inflight backup for
the engine generator.
2. Location
The generator is mounted on the forward face of the APU accessory drive gearbox.
3. Physical Description
Terminal numbering and color-coding is the same as for the engine-driven generators.
Cooling air enters through a duct from outside the APU compartment, passes through the
generator and discharges into the APU compartment. (Duct installation is covered in APU
section).
4. Operation
The APU generator operation is identical to engine generator operation.
5. Maintenance Practices
The generator is attached to the APU accessory drive gearbox with 12 slotted holes over
mounting studs. The APU engine has to be removed prior to removal of the generator.
Removal of the generator is similar to removal of the engine generator.
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AC GENERATOR - APU
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
AC GENERATION
1. Purpose
The purpose of the ac generation system is to provide 115 volt ac, 400 Hz, 3-phase power
for the airplane systems on the ground and in flight.
2. System Description
The system consists of three generators, three generator control units, eight transformers,
four breakers, switches and indicating lights.
GENERAL DESCRIPTION
1. General Component Locations
The components associated with the ac generators are located on the engines (generators,
transformers), APU (generator, transformer), in the electrical equipment compartment
(transformers, breakers), in P6 panel (buses, breakers, generator control units,
transformers) and on P5 panel (switches, lights).
2. General Operation
115 volt ac, 400 Hertz, 3-phase power is supplied by two engine driven generators for normal
inflight operation. A generator, driven by the auxiliary power unit, can supply power for
ground or flight operation.
The control and protection of the generating system is by three identical generator control
units which provide excitation control and fault protection. Power from the generators is
supplied to the buses for distribution through engine and APU generator breakers. Circuit
transformers are used for protection and for power sensing. Power source selection and
indication is from the electrical systems panel.
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AC GENERATION
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COMPONENT FUNCTIONAL DESCRIPTION
AC GENERATOR
Operation
Generator
The generator is rated at 45 kilovolt-amperes 0.95 leading to .75 lagging power factor,
120/208 volts, 400 Hertz alternating current. The unit is without slip-rings,
commutator, or brushes on either the main generator or the exciter. A complete
generator assembly consists of an ac exciter generator, a rotating rectifier and a main
generator. The ac exciter consists of a six-pole stationary dc field, and a rotating
armature. A rotating electromagnetic field causes the output voltage to be induced in
the stationary generator armature. This rotating field is excited by an integral ac
exciter, which has its output converted to dc by rectifiers located in the generator
rotor shaft.
Excitation
The exciter field is supplied with dc power from the voltage regulator. This causes a
3-phase voltage to be developed in the exciter armature which is rectified and fed into
the ac generator rotating field. The field generates the useful ac output voltage in
the ac stator.
Current supplied by the voltage regulator to the field winding provides excitation for
the exciter generator and controls the exciter output to the main generator.
Temperature Control
The thermistor, mounted in the exciter frame, has an inverse temperature resistance
characteristic. The high resistance at low or normal ambient temperature blocks
current flow in one of the parallel wires and causes the overall field resistance to be
about that of the remaining single wire. At the higher temperature resulting from
normal operation, the resistance of each single wire increases to approximately double.
At the same time, the resistance of the two parallel wires at higher temperatures is
approximately equal to that of the single wire at low temperature. Temperature
compensation is thereby provided.
Permanent Magnet Voltage
Permanent magnets are mounted on the exciter generator frame between the six stator
poles. These magnets provide a built-in residual voltage which accrues voltage build-up
and eliminates the need for field flashing or for a starting relay.
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AC GENERATOR
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
GENERATOR CONTROL UNIT
1. Purpose
The purpose of the generator control unit is to provide excitation, control and protection
functions for the generating system.
2. Location
The three generator control units are located on the front of the P6 panel.
3. Physical Description
The GCU is packaged in an ARINC 1/4 short case. A single ARINC hold down hook is attached
directly to the box front face. A circuit breaker is located on the front of the box and
the electrical connector on the rear of the box.
GENERATOR CONTROL UNIT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
CONTROL POWER SUPPLY
Power
Each of the generator control units (GCU) and the bus protection panel (BPP) contains a
transformer-rectifier unit (TR). The TR unit converts 3-phase, ac power to 28 volt dc power
for the control and protection circuits during normal system operation.
If the ac power is not available or the TR unit has failed, the power for the control and
protection circuits is available from the hot battery bus.
With the battery switch in the ON position, relay R41 is energized. The hot battery bus is
connected through the relay to the three GCU's and BPP as a backup power for each of the TR
units.
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CONTROL POWER SUPPLY
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
ENGINE GENERATOR BREAKER
1. Purpose
The two engine generator breakers allow 115 volt ac, 400 Hertz, 3-phase power to be supplied
to the distribution system from the engine generators.
2. Location
The two engine generator breakers are located on either side of a bulkhead attached to the
right outboard side of the nose wheel well.
3. Physical Description
The breaker has a dc coil for closing and tripping. A permanent magnet assists closing and
latches the breaker in the closed position. An internal spring assists opening and holds it
in the open position.
The breaker has 6 main 115 volt ac contacts and 20 auxiliary contacts. The auxiliary
contacts are used to control the positions of APU generator breakers and external power
contactors, and for indicating lights in the flight compartment.
4. Access
Access to the breakers is gained by removing the forward and aft access panels from the
right inner wall in the nose wheel well.
5. Power
Power supply for operation of the breaker is 28 volt dc.
6. Control
The closing and tripping signals for the breakers are from the P5-4 panel, generator control
units and the bus protection panel.
7. Operation
Application of 28 volt dc to close, allows momentary energizing of the coil. When the
breaker closes, the coil deenergizes through the auxiliary contact and the breaker is held
closed by the permanent magnet. Application of 28 volt dc to trip, reverses the polarity of
the coil and the contactor opens assisted by the spring. The coil deenergizes through the
auxiliary contact.
8. Monitor
The breaker position is monitored by the amber BUS OFF and blue GEN OFF BUS lights on the
P5-4 panel.
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ENGINE GENERATOR BREAKER
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COMPONENT FUNCTIONAL DESCRIPTION
DIFFERENTIAL CURRENT TRANSFORMER (ENGINE)
1. Purpose
The purpose of the differential current transformer for the engine generator is to sense the
total current flowing back to the generator.
2. Location
The differential protection neutral current transformer is mounted on the left side of the
engine fan case.
3. Physical Description
Each transformer has 3 windings (one per phase). The generator neutral cables pass through
holes on the transformer marked B1, B2 and B3 facing the generator and are connected to a
common ground on an engine flange.
It is identical to the transformer of the APU generator.
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DIFFERENTIAL CURRENT TRANSFORMER (ENGINE)
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MAINTENANCE TRAINING MANUAL
DIFFERENTIAL CURRENT TRANSFORMER (LOAD)
1. Purpose
The purpose of the differential current transformer for the load (buses) is to sense the
current flowing to the operating systems.
2. Location
Two identical load differential protection current transformer (DPCT) assemblies are located
in the P6 panel.
3. Physical Description
Each current transformer assembly has 6 windings, 2 per phase. One set of windings is used
to compare the load current to the total current furnished by the engine generator and the
other set of windings is used to compare the load current to the total current furnished by
the APU generator.
4. Access
Access to the transformers is by opening P6-11 and P6-12 circuit breaker panels at the
bottom of P6 panel.
DIFFERENTIAL CURRENT TRANSFORMER (LOAD)
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
LINE CURRENT TRANSFORMER ASSEMBLIES
1. Purpose
The purpose of the current transformer assemblies is to provide for the generating system:
Overcurrent Protection
Current Metering
Current Boost
Current Limiting
2. Location
The three line current transformer assemblies (one for each generator) are located on the
right forward ceiling of the electrical equipment compartment.
3. Physical Description
Each assembly contains 12 windings (4 per phase). The generator feeder cables pass through
holes on the transformer marked A1, A2, and A3 facing the generator.
LINE CURRENT TRANSFORMER ASSEMBLYS
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NOTES
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OPERATION
GENERATOR CONTROL
Operation Sequence
Generator Control Unit
Each of the three generator control units (GCUs), located in the P6 panel, contains the
following;
A field power supply (TR unit) which converts 3-phase, ac power from the generator
to a ripple dc voltage for the generator excitation.
A control dc power supply (TR unit) which converts 3-phase, ac power from the
generator to 28 volt dc power for the control and protection circuits for the ac
power system.
A voltage regulator which controls the dc power from the field power supply for
the generator exciter.
A double-coil magnetic-latching generator control relay (GCR), K603, which
connects the output of the field power supply to the generator exciter.
Protection circuits for:
Overvoltage
Undervoltage
Overfrequency
Underfrequency
HV
LV
OF
UF
130
100
430
365
±3
±3
±5
±5
volts
volts
Hertz
Hertz
Overcurrent
OC 170-175 amps
Differential Current
Protection
DP 20-30 amps
Generator Control Relay Operation
The GCR is closed only by placing the generator switch (P5 panel) momentarily to ON.
The GCR is opened by:
Manual: (1)
Generator switch momentarily to OFF.
(2)
Fire handle (P8) pulled ±7 ±2 second time delay).
(3)
CSD disconnect switch activated momentarily (not APU).
Automatic:
(4)
Overvoltage 130 ±3 volts with inverse time delay.
(5)
Undervoltage 100 ±3 volts ±7 ±2 second time delay).
(6)
Overcurrent 170-175 amps with inverse time delay.
(7)
Differential current protection, 20-30A, (25 ms time delay).
The control dc power supply in the generator control unit is backed up by 24 volts dc from
the hot battery bus.
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MAINTENANCE TRAINING MANUAL
OPERATION
Generator Regulator Operation
On initial startup, generator excitation is provided by (3) permanent magnets located in the
excitor stator. The magnets assure that startup voltage is always the same polarity and eliminate
the need to flash the field. The magnetic fields from the magnets induce an AC voltage into the
rotating armature. This AC voltage is changed into a DC voltage by (6) rectifying diodes located
in the generator rotating shaft.
The pulsating DC voltage from the diodes is supplied to the rotating field of the generator and the
resulting magnetic fields induce an AC voltage into the three fixed stator windings. By physically
locating the three stator winding 120 degrees apart three separate AC voltages are generated.
These voltages are called phase A, B, and C and are phased 120 degrees apart.
The AC voltage from the fixed stators is sent to the Voltage Regulator and the Field Power
supply located in the Generator Control Unit (GCU). The purpose of the voltage regulator and
field power supply is to provide excitation to the generator excitor windings and maintain a
generator output of 115 volts, regardless of load, from each phase. Rippled DC is fed from the
field power supply to the DC excitor winding. This current produces more magnetic flux which
adds to the permanent magnet flux.
The excitation from the field power supply is fed through a relay, located in the Generator
Control Unit, called the Generator Control Relay (GCR). This relay can be opened and closed
manually with a switch located on the P5 overhead panel. The relay can also be opened by
generator faults. This prevents faulty electrical power from being applied to the aircraft electrical
system.
If the GCR is open, and the generator is turning only 15 - 20V. AC is produced by the permanent
magnets in the generator. This voltage is called Residual Voltage and is readable on the aircraft
AC voltmeter on the P5 overhead panel. The voltmeter has two scales, the upper scale reads 100 130 volts, and the lower scale reads 0 - 30 volts. By pushing the RESID VOLTS button below
the meter the lower scale is activated allowing the reading. If the button is pushed with the
generator producing full output, GCR closed, the meter can be damaged.
When the generator output reaches 115 volts, the Generator Breaker can be closed and the
generator voltage is supplied to the aircraft buses. Generator output is monitored by fault
detection circuits in the Generator Control Unit (GCU). If the GCU senses faulty electrical power
from the generator, the GCR is opened, and this results in the Generator Breaker opening thus
removing the generator from the aircraft buses.
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GENERATOR CONTROL
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MAINTENANCE TRAINING MANUAL
OPERATION
ANNUNCIATOR PANEL FAULT LIGHTS
Subsystem Sequence
The twelve white malfunction lights in the upper right quadrant of the annunciator panel (4
for each generator) are controlled by double-coil relays inside the generator control units
(GCU).
All malfunction lights can be tested by using the INDICATE/TEST switch to the TEST position.
All lights have 28 volt dc battery bus power available (provided the battery switch is ON).
A ground will be provided within a generator control unit by actuation of the close coil of
the applicable fault light relay. Once the relay has been energized, it is held closed by a
permanent magnetic latch, and can only be released by energizing its reset coil. For the
FF, HV and LV lights, the only manner in which the reset coils can be energized and the
lights extinguished, is by pushing the ERASE button. For the MT light, the only manner in
which its reset coil can be energized is by resetting the applicable generator control relay
(GCR) by manually using that generator's switch to ON. The ERASE button has no effect on MT
lights.
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ANNUNCIATOR PANEL FAULT LIGHTS
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
DC GENERATION
1. Purpose
The 28 volt dc power is used for operation of the airplane systems. The power is obtained
from the battery and from conversion of ac power.
2. System Description
The system consists of a battery, three transformer rectifiers, battery charger, static
inverter, external dc power receptacle, relays, switches and busses.
GENERAL DESCRIPTION
1. General Component Locations
The components associated with the dc power are located in the electronics compartment
(battery, transformer rectifiers, battery charger, static inverter, external dc receptacle),
on P5 panel (switches) and in P6 panel (sensor and relays).
2. General Operation
The 28 volt dc power is obtained primarily from the battery. If ac power is operating, the
three transformer rectifiers (TR No. 1, TR No. 2, TR No. 3) and battery charger supply the
28 volt dc power. The inverter converts battery power to 115 volt ac for the standby
system. The external dc is used for starting APU if battery output is too low. The
distribution system consists of a hot battery bus, battery bus, dc bus No. 1, dc bus No. 2
and standby bus. The control of the distribution system is by three switches located on the
electrical systems panel (P5).
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DC GENERATION
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COMPONENT FUNCTIONAL DESCRIPTION
BATTERY
1. Purpose
The purpose of the battery is to provide dc power to critical airplane systems in the
absence of normal dc supply from the transformer rectifier. It is also used as a backup
power for the ac system control and protection and for starting the APU.
2. Location
The battery is located in the electronic compartment, left side just forward of the E2 rack.
3. Physical Description
The battery is a 20 cell nickel-cadmium unit rated at 36-amp hour. It has a thermostat set
at 58°C (136°F), which is part of battery protection circuit.
4. Monitor.
The output of the battery can be observed on the P5-13 panel.
5. Control
The battery is connected to the dc system by relays operated by switches on the P5 panel.
6. Maintenance Practices.
The nickel-cadmium battery must be serviced at regular intervals determined
conditions. The quantity of water consumed by the battery is determined by
methods and ambient temperatures. Electrolyte level rises when the battery
lowers when the battery discharges. Distilled or demineralized water should
battery cells only when battery is fully charged.
by operating
operating
charges and
be added to
The only battery servicing allowed while the battery is on the airplane is checking
electrolyte level, adding electrolyte, and cleaning spilled electrolyte.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
7. Caution (Cont.)
Keep any metallic objects such as tools, wire brushes, etc., away from exposed top of
battery. Short circuits caused by such objects would be dangerous to personnel.
Remove spilled electrolyte from hands, clothing or other material immediately with water or
3 percent boric acid solution.
Never use hydrometers, droppers, syringes, or other tools on a nickel-cadmium battery which
have been in contact with a lead-acid battery.
Always allow battery to stand idle for 2 to 4 hours with vent caps loosened after being
charged so that all gas may escape before adjusting electrolyte level.
Add only distilled or demineralized water to battery cells. Do not overfill battery.
Remove spilled electrolyte from structure immediately. Wipe up thoroughly with a mild
solution of vinegar and water. Then wash with clear water.
BATTERY
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
BATTERY CHARGER
1. Purpose
The purpose of the battery charger is to restore and maintain the battery at full electrical
potential.
2. Location
The battery charger is located in the electronic compartment on E3-1 shelf.
3. Physical Description
The charger is packaged in an ARINC 3/8 short case. Two ARINC hold-down hooks are attached
directly to the box front face. The electrical connector is on the rear of the box. The
charger is rated at 40 amps with forced air cooling.
4. Power
The battery charger is normally supplied with 115 volt ac power from the ground service bus
and alternately from the main bus No. 2 and supplies dc power to the battery.
5. Control
The battery charger has internal and external controls for the two modes of operation.
5. Operation
The battery charger maintains uniform battery cell voltage levels in two modes of operation.
The high mode, used most of the time, provides rapid charging of the nickel-cadmium
battery, followed by a pulsing charge. Above 16 amperes current, the charger acts as a
normal unregulated transformer-rectifier supply (TR unit). When the battery has sufficient
charge that the current tends to go below 16 amperes, the charging current is reduced
abruptly to zero. The current remains at zero until the battery voltage drops below the
charger voltage. At that time, the charger provides the battery with a pulse of current and
the process repeats itself. The pulsing continues until the mode control circuits change
operation to the low mode, approximately 2 minutes after pulsing begins.
7. Monitor
The output of the battery charger can be observed on the P5-13 panel.
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BATTERY CHARGER
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COMPONENT FUNCTIONAL DESCRIPTION
TRANSFORMER RECTIFIER
1. Purpose
The purpose of the three transformer rectifiers is to convert 115 volt ac, 400 Hz, 3-phase
power to28 volt dc power for use by the airplane's systems.
2. Location
The three transformer rectifier units are located in the electronic compartment on E3-1
shelf.
3. Physical Description
The transformer rectifier is packaged in an ARINC 3/8 short case. A single handle is
attached to the box front face. It is held on the shelf by two hold-down screw-type locks.
The electrical connector is on the rear of the box. The unit is rated at 50 amps with
forced air cooling.
4. Power
The TR No. 1 is supplied with 115 volt ac power from transfer bus No. 1 and supplies 28 volt
dc to dc bus No. 1.
The TR No. 2 is supplied with 115 volt ac power from transfer bus No. 2 and supplies 28 volt
dc to dc bus No. 2.
The TR No. 3 is supplied with 115 volt ac power from main bus No. 2 and supplies 28 volt dc
to battery bus and acts as a backup for TR No. 1 and TR No. 2.
5. Control
The output of TR 3 is connected to the dc system by the battery switch on P5-13 panel.
6. Monitor
The output from each transformer rectifier can be observed on the P5-13 panel.
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TRANSFORMER RECTIFIER
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COMPONENT FUNCTIONAL DESCRIPTION
DC EXTERNAL POWER RECEPTACLE
1. Purpose
The dc external power receptacle is used to connect a 28 volt dc from an external source in
parallel with the airplane battery for starting the APU if the airplane battery voltage is
insufficient.
2. Location
A dc external power receptacle is located below the battery inside the
electrical/electronics compartment.
3. Physical Description
The receptacle consists of two large pins (positive and negative) and a third small pin for
correct orientation of the external connector. The receptacle is enclosed in a housing with
a hinged lid.
4. Control
The connection from the receptacle to the dc system is controlled by a circuit breaker
marked EXTERNAL POWER DC.
5. Operation
The circuit breaker marked BATTERY CHARGER must be open before the application of power, and
the circuit breaker marked EXTERNAL POWER DC must be closed.
The dc external power system is not intended to charge the airplane battery from a dc ground
source. The battery should be charged from the battery charger.
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DC EXTERNAL POWER RECEPTABLE
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COMPONENT FUNCTIONAL DESCRIPTION
STATIC INVERTER
1. Purpose
The purpose of the static inverter is to supply 115 volt ac single phase power to the ac
standby bus during the loss of the normal supply.
2. Location
The static inverter is located in the electronic compartment on E3-1 shelf.
3. Physical Description
The inverter is packaged in an ARINC 2/4 case. Two ARINC hold-down hooks are attached
directly to the box front face. The electrical connector is on the rear of the box. The
inverter is rated at 500 volt amperes with forced air cooling.
4. Power
The inverter is supplied with 28 volt dc power and delivers 115 volt pulsed power to the
standby ac bus.
5. Control
The power supply to the inverter is controlled by the standby power switch on the P5-5
panel.
6. Monitor.
The output of the inverter can be observed on the P5-13 panel.
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STATIC INVERTER
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
DC POWER SYSTEM
1. Purpose
The 28 volt dc power is used for operation, control, and indication of the airplane systems.
2. Location
The 28 volt dc distribution components are located in the electronics compartment, P6 panel
and P5 panel.
3. Control
The control of the dc power is by the battery switch and standby power switch located on P5
panel.
4. Operation
A.
Battery Bus Control
The battery bus is normally powered by the transformer rectifier unit No. 3 (TR3),
through battery bus relay R355 which is energized with the battery switch ON. The
output of TR3 powers relay R2 bat transfer, which relaxes bat bus auto relay R1 allowing
TR3 output to power the battery bus through relays R1 and R326 (relay R326 is normally
relaxed).
If the output of TR3 is not available, relay R2 relaxes, relay R1 energizes and the
battery bus is powered from the hot battery bus.
If the standby power switch is placed to the BAT position, relay R326 energizes,
disconnecting TR3 and connecting the hot battery bus to the battery bus.
B.
Battery Charger Control
The 115 volt ac 3-phase source of power for the battery charger is normally supplied
from the ground service bus, which in turn is normally powered by the generator bus No.
1.
If the generator No. 1 trips off, GB1, APU GB1 and EPC1 will be open. The same power
from GCU2 which energizes the alternate coil of transfer relay No. 1 is used to
energize the battery charger transfer relay R89. This relay transfers the battery
charger to its alternate source of power, main bus No. 2.
Between relay R89 and the battery charger, the 3-phase ac power passes through relaxed
battery overheat relay R325 and relaxed APU start interlock relay R39. Relay R39 is
connected in parallel with the APU start relay R5, which is energized while the APU
start motor is cranking and relaxes automatically when the APU reaches starter cutout
speed.
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COMPONENT FUNCTIONAL DESCRIPTION
B.
Battery Charger Control (Cont.)
The battery overheat relay R325 energizes whenever the battery internal temperature
reaches 135ºF. With the relay energized, the ac power supply to the charger is
interrupted.
The battery charger maintains uniform battery cell voltage levels in two modes of
operation.
In addition to the internal control of the mode control relay, the ground for the relay
passes through a loop interconnecting three external control functions. If any one of
the three breaks the ground loop, the charger reverts to the low charge mode.
The R1 aux relay R327 energizing due to loss of the normal input (TR3) to the
battery bus. (High mode pulsing by the charger would result in pulsing of the
flight compartment indicating lights, which is undesirable).
The R328 energizing by the action of the standby power switch to the BAT position.
(High mode pulsing by the charger would result in pulsing of the flight
compartment lights, which is undesirable).
The refueling power select relay R10 energizing with external ac power connected
to the airplane. (Faulty regulation of ground power could damage the battery if
the charger is operating in the high mode as an unregulated transformerrectifier).
C.
Power Distribution
The normal dc power supply system consists of three interchangeable 50 ampere
transformer-rectifier (TR) units.
Transformer-rectifier No. 1 (TR1) is powered by 3ø 115 volt ac from transfer bus No. 1
and delivers 28 volt dc to dc bus No. 1. TR2 is powered by 3φ 115 volt ac from transfer
bus No. 2 and delivers 28 volt dc to dc bus No. 2. TR3 is powered by 3φ 115 volt ac
from main bus No. 2 and delivers 28 volt dc through battery bus relay R355, bat bus
auto relay R1 and bat bus manual relay R326 to the battery bus. The battery switch, on
P5 panel, must be ON. If TR3 has no output, the battery bus is connected to the hot
battery bus through R326; the battery switch must again be ON. TR3 is also connected
through a diode to dc bus No. 2, and through normally energized TR3 disconnect relay R9
to dc bus 1. The loss of a single TR unit will not result in the loss of a dc bus. R9
can be de-energized by the bus transfer switch placed to OFF.
5. Monitor
The monitoring of the dc power is by the Annunciator Panel M238 and Electrical Power Systems
Test Module M400 in conjunction with meters on P5
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DC POWER SYSTEM
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
STANDBY POWER DISTRIBUTION
1. Purpose
The purpose of the standby power distribution is to provide 115 volt ac and 28 volt dc power
to essential system during absence of normal supply.
2. Location
The standby power components are located in the electronics compartment, P6 panel and P5
panel.
3. Control
The control of the standby power is by switches and relays.
4. Operation
A.
DC System
The standby dc system is 28 volt dc bus normally powered from dc bus No. 1 through
energized standby power transfer auto relay R37 and relaxed standby power manual relay
R328. The standby power switch being in the AUTO position. On the ground standby
power off relay R356 is energized and the static inverter is prevented from operating
unless ac power is available to transfer bus No. 1 through the R3 transfer relay No. 2
in the NORMAL position.
The standby bus can also be powered from the battery bus by placing the standby power
switch to the BAT position, thus energizing relay R328.
With the standby power switch placed to the OFF position no signal will reach the
static inverter to turn it on.
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COMPONENT FUNCTIONAL DESCRIPTION
B.
AC System
The standby ac system is a single-phase, 115 volt ac bus normally powered from transfer
bus No.1 through energized standby power transfer auto relay R37 and relaxed standby power
manual relay R328. Relay R37 is powered from dc bus No. 1 and is grounded through
transistor Q1 and standby power switch in AUTO position. The output of transfer bus No. 1
must be greater than 100 volt ac to the voltage sensor for bias of transistor Q1.
With standby ac bus energized, standby power ind light relay R330 is energized and the
STANDBY POWER OFF amber light on P5 panel is extinguished.
When the output of transfer bus No. 1 decreases to below 100 volt ac or with loss of power
on dc bus No. 1, transistor Q1 is unbiased and relay R37 relaxes. The standby inverter,
powered from the battery bus, will now supply the ac power to the standby bus through
relaxed relays R37 and R328.
The standby bus can also be powered from the standby inverter by placing the standby power
switch to the BAT position, thus energizing relay R328.
With the standby power switch placed to the OFF position, relays R37 and R328 are relaxed and
the standby inverter receives no power from the battery bus and the standby bus will not
be powered.
Whenever the standby bus is not powered, relay R330 is relaxed and the STANDBY POWER
OFF light is illuminated if battery switch is ON.
5. Monitor
The monitoring of the standby ac and dc power is by the Annunciator Panel M238, and
Electrical Power System Test Module M400 and meters on P5 panel.
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STANDBY POWER DISTRIBUTION
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INTRODUCTION
AC EXTERNAL POWER
1. Purpose
The 115 volt ac, 400 Hz, 3-phase power from an external source (ground power unit) can be
supplied to the airplane for systems use.
2. System Description
The system consists of an external power receptacle, bus protection panel for control and
protection functions, two external power contractors, relays, switches and indicating
lights.
GENERAL DESCRIPTION
1. General Component Locations
The components associated with the external power are located in the P6 panel, (busses, BPP
and ground service relay), on P5 panel (switch, light), P13 panel (switch), nose wheel well
area (EPC No. 1 and EPC No. 2) and right side of the forward fuselage (receptacle).
2. General Operation
The ac external power can be supplied to the busses through closed external power contactors
using the switch on P5 panel. For servicing of the airplane, using the ground service
switch energizes the ground service bus. The external bus power can be used for fueling of
the airplane.
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AC EXTERNAL POWER
69
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NOTES
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BUS PROTECTION PANEL
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COMPONENT FUNCTIONAL DESCRIPTION
AC POWER DISTRIBUTION
1. Purpose
The purpose of the ac power distribution is to provide 115 volt ac, 400 Hz, 3-phase power
from various sources to the airplane operating systems.
2. Location
The components of the distribution system are located in the P6 panel and P5 panel.
3. Control
The control of the ac distribution system is by switches and relays.
4. Operation
The three principles of operation of the electrical power system area:
No paralleling of ac sources of power.
The source of power switched onto or entering the system takes priority and will trip
off the existing source.
A source of power does not enter the system automatically when it reaches proper voltage
and frequency. It must be manually switched on by a switch.
The two generator buses can be powered by the ac power sources (engine generators, APU
generator, external power) through the six power breakers.
The power distribution is from the generator buses to the transfer buses, ground service
bus, dc system, electronics system and standby system.
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AC POWER DISTRIBUTION
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
GALLEY SERVICE POWER
1. Power
Three-phase, 115 volt, 400 Hertz, ac power is supplied to gallery units for ovens, coffee
makers, or other electrically powered units.
2. Control
The control switch for gallery electrical power, located on the P-5 panel, provides a ground
for the galley power relays. 28 volts dc for the relays comes from opposite generator
control units. Loss of either generator will automatically cause a loss of power to all
galleys.
GALLEY SERVICE POWER
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MAINTENANCE PRACTICES
AC AMMETERS
Test
The three ac ammeters on the P5 overhead panel for the two engine generators and the APU
generator normally read phase B line current. Each of the ammeters senses current using one
of the metering transformer windings in its line current transformer assembly in the ceiling
of the electrical/electronics compartment.
The ammeters can read phase A, B or C current by selecting positions of the switch S2 on the
Power System Test Module M400 in the P6 panel.
For each ammeter, the connection from the line current transformer windings to the ammeter
is made through contacts of three relays inside the M400 module. The relative positions of
the relays are controlled by switch S2.
If the switch is in position A or D, relays K1 and K2 are energized and the phase A windings
are connected to all three ammeters and all other current transformer windings are in an off
status by being shorted to themselves.
If the switch is in position B, E, G of H, all three relays are relaxed and the phase B
windings are connected to all three ammeters. Position B is normal so that all electrical
power ac meters on the overhead panel read phase B voltage, frequency and current. Also,
the relays K1, K2 and K3 are not held energized unnecessarily.
If the switch is in position C or F, relays K2 and K3 are energized and the phase C windings
are connected to all three ammeters.
For the APU generator, the line current transformer also provides overcurrent protection by
positioning off the galley switch. This would trip galley load before the APU GCU
overcurrent protection trips the APU field (GCR) and APU generator breaker.
EFIS Difference
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AC AMMETERS
EFIS Difference
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GENERAL DESCRIPTION
ELECTRICAL SYSTEM PANELS
General Component Locations
The operation and indication of the various systems is from the overhead panel (P5) in the
flight compartment. Three sections of the panel are for the electrical power system.
AC and DC Meter Panel (P5-13)
Constant Speed Drive and Standby Power Panel (P5-5)
Generator Ammeter and Bus Switching Panel (P5-4)
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ELECTRICAL SYSTEM PANELS
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GENERAL DESCRIPTION
LOAD CONTROL CENTER
General Component Locations
The circuit breaker panels behind the first officer and the captain are the main load
control centers and contain the ac and dc buses of the electrical power system.
Panels P6-1, P6-2, P6-3, P6-4, P18-2, P18-3, and P18-4 are hinged at the bottom. Panels P611,P6-12, P18-1 and P18-5 are hinged at the side.
The P6 panel also contains the three interchangeable generator control units (GCUs) and the
bus protection panel (BPP).
To the left of the generator control units are two receptacles that can be used for
connecting cleaning or test equipment.
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LOAD CONTROL CENTER
81
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NOTES
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P6 PANEL INTERIOR
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RELAY ASSEMBLY (P6-2 PANEL)
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RELAY ASSEMBLY (P6-2 PANEL)
B737 - EFIS AIRPLANE
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ELECTRICAL POWER DISTRIBUTION
86
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ELECTRICAL POWER DISTRIBUTION
B737 - EFIS AIRPLANE
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MAINTENANCE PRACTICES
AC AND DC METERS
AC Volts and Frequency
The ac volts and frequency may be read on the P5 ac volt meter and frequency indicator
for standby power (φA), ground power, generator No. 1, APU generator, generator No. 2
(φB) and inverter (φA) by selecting each of these with the selector switch. When the
switch is in the TEST position, 24 additional test points may be selected from
combinations of the two M400 selector switches S1 positions 1, 2, 3 and S2 positions A,
B, C, ... H. These 24 test points may also be selected for reading ac volts and
frequency on portable meters connected to the M400 Test Module ac jacks.
DC Volts
The dc volts can be read on the P5 overhead panel dc voltmeter for standby bus,
battery bus, battery, and each of the three transformer-rectifiers. By using the
overhead panel selector switch in the TEST position, 32 additional test points may be
selected from combinations of the two M400 selector switches S1 positions 2, 3, ... 8)
and S2 positions A, B, C, D. These 32 test points may also be selected for reading dc
volts on a portable dc voltmeter connected to the M400 Test Module dc jacks.
AC AND DC METER PANEL (P5-13)
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ELECTRICAL POWER ANNUNCIATOR PANEL M238
The Electrical Power Annunciator Panel M238 is located in a well on the aft side of the P6
panel in the flight compartment entrance way. It is used for the indication of power on the
individual ac and dc buses and indication of a fault which tripped the generator control
relay GCR. Under normal operation, with the shield closed, no lights will be visible on the
annunciator panel.
The panel is divided into three sections: ac bus lights (amber neon), dc bus lights (white), and
malfunction lights (white).
The eleven neon lights in the lower right hand quadrant are illuminated when phase A and C
of their respective ac buses are powered. Each lights is connected to a circuit breaker on its
respective bus and all lights are connected to a common ground.
The six DC lights in the upper left hand quadrant are connected to a circuit breaker on their
respective buses. Ground for the lights is provided through a toggle switch located on the
panel. To test the DC buses the toggle switch must be held in the INDICATE position. The
EXT PWR TR light, lower left dc light, will only operate if external power is connected to
the aircraft and the power is of usable quality.
The six malfunction lights on the upper right hand quadrant provide indications of faults
that have tripped the generator control relay GCR for the respective generator. One light
in each series, MANUAL TRIP (MT) when illuminated indicates that the pilot has tripped
the GCR with the generator control switch of with the CSD disconnect switch.
ELECTRICAL POWER ANNUNCIATOR PANEL M238
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MAINTENANCE PRACTICES
POWER SYSTEM TEST MODULE M400
The Power System Test Module M400 is located in a well on the aft side of the P6 panel in
the flight compartment companion way and is used to select DC volts, and AC volts and
frequency at various test points in the electrical system. The selection is made by rotary
switches S1 and S2. Readings can be displayed on the P5-13 panel in the flight compartment,
or on separate, portable meters connected to the M400 ac od dc jacks.
POWER SYSTEM TEST MODULE M400
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MAINTENANCE TRAINING MANUAL
NOTES:
1. TO READ: (A) DC VOLTS SET P5 DC SELECTOR SWITCH TO TEST AND ADJUST
M400 S1 AND S2
(B) GEN AMPS AJDUST S2 FOR PHASE
(C) AC VOLTS AND FREQUENCY SET P5 AC SELECT SWITCH TO
TEST AND ADJSUT M400 S1 AND S2
2. ALL READINGS CAN BE MADE ON THE P5 METERS OR PRECISION METERS
CONNECTED TO THE AC OR DC JACKS ON THE M400.
3. S2 IS NORMALLY LEFT IN THE B POSITION CONNECTING ALL 3 GENERATOR
AMMETERS TO PHASE B AND LEAVING THE M400 SEELCT RELAYS RELAXED.
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NOTES
92
ARINC 429
and
Abbreviations
1
MAINTENANCE TRAINING MANUAL
NOTES
2
MAINTENANCE TRAINING MANUAL
INTRODUCTION
DITS - DATA TRANSMISSION CHARACTERISTICS
1. Development of Data Buses
The gradual changes from primarily analog data transmission to primarily digital
transmission methods was caused by the increasing use of digital processing within the
system LRU's, and a need to reduce the amount of aircraft wiring.
With the development of digital techniques and the advent of microcomputers more systems
LRU's are using digital techniques to implement new functions or functions previously
performed by analog circuits.
2. MARK 33 DITS "Basic Philosophy"
A.
Description
With the advent of digital technology serial digital data buses evolved to replace
their analog counterparts. The MARK 33 DITS specification describes a method of
transmitting information over a shielded, twisted pair of wires.
Bi-directional data flow is not permitted. This means that a separate data bus (a
shielded, twisted pair of wires) is required for each direction when data is to flow
between 2 system elements.
B.
Data Flow
In the example, the DITS interface is represented by boxes X, Y, and Z. Data is
transmitted in one direction on a pair of shielded wires. Box "X" supplies data from
transmitting port "A" to receiving ports "A" on boxes "Y" and "Z". These decode only
the data required by the individual boxes. Data is also transmitted from port "C" of
box "Z" to receiving port "D" of box X.
Each data bus will originate from 1 transmitter and go to a maximum of 20 receivers.
Component parts are designated as either transmit or receive exclusively.
3
MAINTENANCE TRAINING MANUAL
INTRODUCTION
C.
Data Transfer - General Discussion
NOTE:
Detailed coverage of this "general discussion" - will follow.
Data for transmission is encoded in either two's complement fractional binary notation
(BNR) or binary coded decimal notation (BCD). In addition, alpha-numeric data is
transmitted encoded per ISO alphabet #5. This data is supplied from source systems at
rates sufficiently high to ensure small incremental value changes between updates.
Transmission is made "open loop", i.e. the receiving system is not required to inform
data sources that information has been received.
A "parity bit" is transmitted as part of each data word to permit simple error checks
by the receiving system. Together, with "data reasonableness" checks, also performed by
the receiving system, the error checks may be used to prevent the display, or use of an
erroneous or suspect word.
The high integrity of the shielded, twisted wire "transmission medium" ensures that
such missing bits are few.
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DITS - DATA TRANSMISSION CHARACTERISTICS
5
MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
DIGITAL WORD AND LABEL
1. Digital Word - Information Element
Digital data is transmitted by sending electronic pulses in a series over the digital buses.
An individual pulse is referred to as a "Bit".
The basic information element is a digital word containing 32 bits. There are several
application groups for such words, which will be addressed later.
2. Label - Information Identifier
The significance of a Bit within a word, depends on the Bit position. The first 8 bits of
each word are used as labels. The label identifies the information contained within the
word (i.e. Cabin Component Temperature).
DIGITAL WORD AND LABEL
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
SOURCE/DESTINATION IDENTIFIER AND DATA FIELD
1. Source/Destination Identifier (SDI)
The Bits 9 and 10 comprise the source/destination identifier, or SDI. The SDI function is
used when it is necessary to indicate the source of information, or when the information is
directed to a specific location. As an example, when specific words need to be directed to
a specific system of a multi-system installation, and when the source system of a multi
system installation needs to be recognizable from the word content.
2. Data Field
The data field, bits 11 through 28, or 29, depending on word type, contains the specific
data assigned to a label.
SOURCE/DESTINATION IDENTIFIER AND DATA FIELD
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
SIGN STATUS MATRIX AND PARITY
1. Sign Status Matrix (SSM)
The Sign Status Matrix code, or SSM, consists of bits 29 or 30 through 31, depending on word
type. The SSM code identifies the characteristics of a word (plus, minus, north, south,
etc.) and its status (no computed data, test, etc.)
2. PARITY
The last bit, number 32, is a parity bit used for checking transmission efficiency.
SIGN STATUS MATRIX AND PARITY
8
MAINTENANCE TRAINING MANUAL
Abbreviations
9
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
A
AAM
A/C
AC
ACARS
ACK
ACMS
ACP
ACQ
ACT
A/D
ADC
ADF
AFC(S)
AFDS
AGC
ALPHA
ALT
AM
AMP
ANCDU
ANN
ANS
ANT
AOA
AOC
A/P
APP
APU
ARINC
ARPT
ARR
A/S
ASA
Amber
Autopilot Actuator Monitor
Aircraft
Alternating Current
Aircraft Communications Addressing and Reporting System
Acknowledge
Aircraft Condition Monitoring System
Audio Control Panel
Acquire
Actuator
Analog to Digital
Air Data Computer
Automatic Direction Finder
Automatic Flight Controls (System)
Autopilot Flight Director System
Automatic Gain Control
Angle of Attack
Altitude or Alternate
Amplitude Modulation
Amplifier
Alternate Navigation Control Display Unit
Annunciator
Alternate Navigation System
Antenna
Angle of Attack
Approach on Course
Autopilot
Approach
Auxiliary Power Unit
Aeronautical Radio Incorporated
Airport
Arrival
Airspeed
Autoflight Status Annunciator
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MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
A/T
ATA
ATC
ATCRBS
ATE
ATIS
A/T LIM
ATR
ATT
ATTEND
AUTO
AUX
B
BARO
BAT
BBL
B/CRS
BCD
BCF
BFO
BITE
BNR
BOT
BOV
BRT
C
CAA
CAL
CAPT
CAS
CAT
CB
CCW
CDU
CDX
CG
Autothrottle
Air Transport Association
Air Traffic Control
Air Traffic Control Radar Beacon System
Automatic Test Equipment
Automatic Terminal Information Service
Autothrottle Limit
Austin Trumbull Radio
Attitude (Mode)
Attendant
Automatic
Auxiliary
Blue
Barometric
Battery
Body Buttock Line
Backcourse
Binary Coded Decimal
BromoChlorodiFluoromethane (Halon)
Beat Frequency Oscillator
Built-in Test Equipment
Binary Numerical Reference
Beginning of Tape
Bar Out of View
Brightness
Centigrade
Civil Aviation Authority
Calibrate
Captain
Computed Airspeed
Category Type of Landing
Circuit Breaker
Counterclockwise
Control Display Unit
Control Differential Transformer
Center of Gravity
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MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
CH
CHR
CI
CKT(S)
CLB
CLR
CMD
C/O
COAX
COMM
CON
CONT
COS
CP
CPU
CRS
CRT
CRZ
CT
CTR
CW
CWS
D
D/A
DAA
DADC
DACS
db
DBL
DC
DEL
DEMUX
DEP
DEP ARR
DES
Channel
Chronograph
Cost Index
Circuit(s)
Climb
Clear
Command
Change Over
Coaxial
Communication
Continuous or Maximum Continuous Thrust
Continued
Cosine
Control Panel
Central Processing Unit
Course
Cathode Ray Tube
Cruise
Control Transformer
Center
Clockwise
Control Wheel Steering
Day
Digital to Analog
Digital Analog Adapter
Digital Air Data Computer
Digital Audio Control System
Decibel
Data Base Loader
Direct Current
Delete
Demultiplexer
Departure
Departure Arrival
Descent
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MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
DEST
DEV
DFCS
DFDAU
DFDR(S)
DH
DIR INTC
DLA
DLK
DMA
DME
DN
DSPY
DTG
DU
E
EADI
ECON
E/D
E/E
EEPROM
EFI
EFIS
EGT
EHSI
EIS
ELEV
ELEX
ELT
EMI
ENG
ENGA
ENT
EOT
Destination
Deviation
Digital Flight Control System
Digital Flight Data Acquisition Unit
Digital Flight Data Recorder (System)
Decision Height
Direct Intercept
Delay
Downlink
Direct Memory Access
Distance Measuring Equipment
Down
Display
Distance to Go
Display Unit
East
Electronic Attitude Director Indicator
Economy
End of Descent
Electronic Equipment (compartment)
Electric Erasable Programmable Read Only Memo
Electronic Flight Instruments
Electronic Flight Instrument System
Exhaust Gas Temperature
Electronic Horizontal Situation Indicator
Engine Instrument System
Elevation
Electronics
Emergency Locator Transmitter
Electromagnetic Interference
Engine
Engage
Enter
End of Tape
24
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
EPROM
EQUIP
ESDS
ET
ETA
ETE
EXEC
EXP
F
FAA
FCC
FD
FF
FGN
FL
FLT
FLT CONT
FM
FMA
FMC
FMCS
FMS
FO
FOB
FPA
FPM
FREQ
F/S
FS
FT
FWD
GA
GEN
GMT
GND
GPWC
Erasable Programmable Read Only Memory
Equipment
Electro Static Discharge Sensitive
Elapsed Time
Estimated Time of Arrival
Estimated Time En route
Execute
Expanded
Fahrenheit
Federal Aviation Administration
Flight Control Computer
Flight Director
Fuel Flow or Flip Flop
Foreign (off side) opposite
Flight Level
Flight
Flight Controls
Frequency Modulation
Flight Mode Annunciator
Flight Management Computer
Flight Management Computer System
Flight Management System
First Officer
Fuel On Board
Flight Path Angle
Feet Per Minute
Frequency
Fast Slow
Fast Slew
Feet
Forward
Go Around
Generator
Greenwich Mean Time
Ground
Ground Proximity Warning Computer
25
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
GPWS
G/S
GS
GSE
GW
HDG
HDG HOLD
HDG SEL
HDPH
HEX
HF
HG
HLD
HR(S)
HRZ
HYD
Hz
IAS
I/C
ICAO
IDENT
IDU
IF
IFSAU
ILS
IN
INBD
IND
INFLT
INIT REF
INOP
INPH
INST
INSTR
INT
INTLK
Ground Proximity Warning System
Glide Slope
Ground Speed
Ground Support Equipment
Gross Weight
Heading
Heading Hold
Heading Select
Headphone
Hexadecimal
High Frequency
Mercury
Hold
Hour(s)
Horizontal
Hydraulic
Hertz
Indicated Airspeed
Interphone
International Civil Aviation Organization
Identification
Interactive Display Unit
Intermediate Frequency
Integrated Flight System Accessory Unit
Instrument Landing System
Inches
Inboard
Indicator
Inflight
Initialization Reference
Inoperative
Interphone
Installation
Instrument
Interphone
Interlock
26
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
INV
I/O
IRS
IRU
ISA
ISDU
ISO
K
Kg
KHz
KIAS
KT(S)
L
LAT
LB(S)
LCD
LCL
LED
LH
LIM
LNAV
LOC
LON
LRRA
LRU
LSB
LSK
LT
LVDT
LVL CHG
M
M
MA
MAINT
Inverter
Input Output
Inertial Reference System
Inertial Reference Unit
International Standard Atmosphere
Inertial System Display Unit
International Standards Organization
Kilo (one thousand)
Kilogram (1,000 grams)
KiloHertz (1,000 Hertz)
Knots Indicated Airspeed
Knot(s)
Left
Latitude
Pound(s)
Liquid Crystal Display
Local (on side)
Light Emitting Diode
Left-hand
Limit or Laser Intensity Monitor
Lateral Navigation
Localizer
Longitude
Low Range Radio Altimeter
Line Replaceable Unit
Lower Side Band
Line Select Key
Light or Left
Linear Variable Differential Transformer
Level Change
Master or Meter(s) or Month
Minus (sign) or Mach
Master
Maintenance
27
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
MAN
MASI
MAX
MB
MCDU
MCP
MCU
MD&T
MET
MHz
MIC
MIN
MISC
MKR
MLG
MM
Mmo
MOM
MSG
MSU
MTA
MTS
MU
MUX
mW
N
N1
N2
NA
NAV
NC
NCD
NDB
NEG
NM
NORM
NSS
OAT
Manual
Mach Airspeed Indicator
Maximum
Millibars
Multipurpose Control Display Unit
Mode Control Panel
Modular Concept Unit
Master Dim and Test
Metric
MegaHertz (1,000,000 Hertz)
Microphone
Minimum
Miscellaneous
Marker
Main Landing Gear
Maintenance Manual
Mach maximum operating
Momentarily
Message
Mode Select Unit
Mach Trim Actuator
Mach Trim System
Management Unit
Multiplexer
MilliWatts (one thousandth of a watt)
North
Low Pressure Rotor Speed
High Pressure Rotor Speed
Not Applicable
Navigation
Not Connected
No Computed Data
Navigation Data Base
Negative
Nautical Miles
Normal
Neutral Shift Sensor
Outside Air Temperature
28
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
OBS
OC
OD
OOOI
OP
OSS
OVHD
OVHT
PA
PASS
PB
PCU
PED
PES
PIREP
PLA
PMC
PNL
POS
POT
PP
PPOS
PPS
PREV PAGE
PRI
PROG
PROV
P/RST
PRT
P/S
PSI
PSU
PTH
Observer
On Course
Out of Detent
Out Off On In
Oil Pressure
Over Station Sensor
Overhead
Overheat
Passenger Address
Passenger
Push Button
Power Control Unit
Primary Engine Display
Passenger Entertainment System
Pilot Report
Power Lever Angle
Power Management Control
Panel
Position
Potentiometer
Program Pin
Present Position
Pulses Per Second
Previous Page
Primary
Progress or Program
Provisional
Push to Reset
Printer
Pitot Static
Pounds per Square Inch
Passenger Service Unit
Path
29
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
PTT
Pwr
QAR
QRH
R
RA
RAM
RAT
R-CLB
RCP
RCWS
RDDMI
REC
REF
REQ
RES
RETF
REU
REV
RF
RH
RMI
RNAV
ROM
RPM
RR
RST
R/T
RTA
RTE
R-TO
RVR
RWY
S
SAT
Press To Talk
Power
Quick Access Recorder
Quick Reference Handbook
Red or Right
Radio Altitude or Resolution Advisory
Random Access Memory
Ram Air Temperature
Reduced Climb
Radio Communication Panel
Roll Control Wheel Steering
Radio Direction Distance Magnetic Indicator
Receiver
Reference
Requisition
Resolver
Return To Field
Remote Electronics Unit
Reverse or Revision
Radio Frequency
Right-hand
Radio Magnetic Indicator
Radio Navigation
Read Only Memory
Revolutions per Minute
Rising Runway
Reset
Receiver Transmitter
Required Time of Arrival
Route
Reduced Takeoff
Runway Visual Range
Runway
South
Static Air Temperature
30
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
SB
S/C
SCC
SCR
SDI
SEC
SEL
SEL SPD
SELCAL
SENS
SG
SIDS
SIN
SPD
SPKR
SPM
SQL
SS
SSB
SSEC
SSM
SSR
SSSV
STA
STAB
STARS
STBY
STC
STS
SVC
SYS
SW
SWC
T
TA
TAI
TAS
TAT
Side Band
Step Climb
Station Control Card
Silicon Controlled Rectifier
Source Destination Identifier
Seconds
Selected or Selection
Selected Speed
Selective Calling
Sensitivity
Symbol Generator
Standard Instrument Departure Sequence
Sine
Speed
Speaker
Surface Position Monitor
Squelch
Slow Slew
Single Side Band
Static Source Error Correction
Sign Status Matrix
Secondary Surveillance Radar
Solid State Stored Voice
Station
Stabilizer
Standard Terminal Arrivals
Standby
Sensitivity Timing Control
Status or Speed Trim System
Service
System
Switch
Stall Warning Computer
Time
Traffic Advisory
Thermal Anti-ice
True Airspeed
Total Air Temperature
31
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
T/C
TCAS
T/D
TEMP
TGT
THR HOLD
TLA
TMA
TO
TO/GA
TPR
T/R
TR
TRK
TT
TURB
TX
u
uA
UHF
ULB
uP
USB
usec
UTC
V
V1
V2
VAR
VERT
VHF
Vmo
VNAV
VOR
Top of Climb
Traffic Alert and Collision Avoidance System
Top of Descent
Temperature
Target
Throttle Hold
Thrust Level Angle
Thrust Mode Annunciator
Takeoff
Takeoff/Go Around
Transponder
Thrust Reverser
Transformer Rectifier
Track
Torque Tube
Turbulence
Transmit or Synchro Transmitter
Micro (1 millionth)
Micro Amp
Ultra High Frequency
Underwater Locating Beacon
MicroProcessor
Upper Side Band
Microsecond
Coordinated Universal Time
Voltage or Voice
Takeoff Decision Speed
Target Speed
Variable
Vertical
Very High Frequency
Velocity maximum operating
Vertical Navigation
Very High Frequency Omni Range
32
MAINTENANCE TRAINING MANUAL
ABBREVIATIONS
Abbreviation
Meaning
VOR LOC
VOX
VR
VREF
V/S
VSI
VSWR
VTK
W
WD
WL
WPT
WT
WX
WXR
X
XFER
XFMR
XMIT
XMTR
XPDR
XTK
Y/D
YR
Z
ZFW
VOR Localizer
Voice Mode
Rotation Speed
Reference Speed
Vertical Speed
Vertical Speed Indicator
Voltage Standing Wave Ratio
Vertical Track
White or West or Watts
Wind Direction
Water Line
Way Point
Weight
Weather
Weather Radar
Cross
Transfer
Transformer
Transmit
Transmitter
Transponder
Cross Track
Yaw Damper
Year
Coordinated Universal Time (Zulu)
Zero Fuel Weight
33
MAINTENANCE TRAINING MANUAL
BITE Abbreviations
34
MAINTENANCE TRAINING MANUAL
SORTED BITE ABBREV
A/C INIT
A/C PWR
A/P
A/P DISC SW
A/S
A/T
A/T CMPTR
AC
ACC
ACT
ADI
AFCS
AFDS
AIL
AIL ACT D/W
AIL D/P SW
ALT
ALT ACQ
ANCDU
ANN
ANS
APPS PGM
AT
ATT
AUTH
AUX
BARO
BARO COR SL
BAT
C/B
C/O
CAA
CAT
CCW
CHAN
CMD
CMDS
COL ACTD MODULE
COND
CONN
CONT
CP
CPT
CRC
CRS
CW
AIRCRAFT INITIALIZATION
AIRCRAFT POWER
AUTOPILOT
AUTOPILOT DISCONNECT SWITCH
AIRSPEED
AUTOTHROTTLE
AUTOTHROTTLE COMPUTER
ALTERNATING CURRENT
ACCESSORY
ACTUATOR
ATTITUDE DIRECTOR INDICATOR
AUTOMATIC FLIGHT CONTROL SYSTEM
AUTOPILOT FLIGHT DIRECTOR SYSTEM
AILERON
AILERON ACTUATOR DETENT WRAP
AILERON DETENT PRESSURE SWITCH
ALTITUDE
ALTITUDE ACQUIRE
ALTERNATE NAVIGATION CONTROL DISPLAY UNIT
ANNUNCIATOR
ALTERNATE NAVIGATION SYSTEM
APPLICATIONS PROGRAM
AUTOTHROTTLE
ATTEMPT OR ATTITUDE
AUTHORITY
AUXILIARY
BAROMETRIC
BAROMETRIC CORRECTION SELECT
BATTERY
CIRCUIT BREAKER
CHANGE OVER (SWITCH ON MODE CONTROL PANEL)
CIVIL AVIATION AUTHORITY
CATEGORY
COUNTER CLOCKWISE
CHANNEL
COMMAND
COMMANDS
COLUMN ACTUATED MODULE
CONDITION
CONNECTION
CONTROL
CONTROL PANEL
CAPTAIN
CYCLIC REDUNDANCY CHECK
COURSE
CLOCKWISE
1
MAINTENANCE TRAINING MANUAL
SORTED BITE ABBREV
CWS
CWS OV/RD
D/S
DAA
DADC
DEG
DEV
DFCS
DIMEN
DISC
DISINWD
DME
DSPL
E AIL LM
EADI
EFIS
EIS
EIS-PRI
ELE ACT D W
ELE D/P SW
ELEC AIL LIM
ELEC TRM LK
ELECT
ELEV
ENG
ENG SOL WPH
ENG SOL WPL
F LIM ITLK
F/D
F/O
FAA
FCC
FCC COMP
FGN
FGN AIL D/P
FL PLCRD
FLT
FLT CONT
FLT INST ACC BOX
FMA
FMC
FMS
FMS WARN ANN
FRC LIM C/E
FREQ
FS2
CONTROL WHEEL STEERING
CONTROL WHEEL STEERING OVERRIDE
DEGREES PER SECOND
DIGITAL ANALOG ADAPTOR
DIGITAL AIR DATA COMPUTER
DEGREE
DEVIATION
DIGITAL FLIGHT CONTROL SYSTEM
DIMENSION
DISCONNECT
DISCRETE INPUT WORD
DISTANCE MEASURING EQUIPMENT
DISPLAY
ELECTRONIC AILERON LIMITER
ELECTRONIC ATTITUDE DIRECTOR INDICATOR
ELECTRONIC FLIGHT INSTRUMENT SYSTEM
ENGINE INDICATING SYSTEM
ENGINE INDICATING SYSTEM - PRIMARY
ELEVATOR ACTUATOR DETENT WRAP
ELEVATOR DETENT PRESSURE SWITCH
ELECTRONIC AILERON LIMITER
MAIN ELECTRIC STABILIZER TRIM SWITCH
ELECTRONIC OR ELECTRICAL
ELEVATOR
ENGAGE
ENGAGE SOLENOID WRAP HIGH
ENGAGE SOLENOID WRAP LOW
AILERON FORCE LIMITER INTERLOCK
FLIGHT DIRECTOR
FIRST OFFICER
FEDERAL AVIATION ADMINISTRATION
FLIGHT CONTROL COMPUTER
CONFIGURATION PIN
FOREIGN
FOREIGN AILERON DETENT PRESSURE
HIGH FLAP PLACARD OPTION
FLIGHT
FLIGHT CONTROL
FLIGHT INSTRUMENT ACCESSORY BOX
FLIGHT MODE ANNUNCIATOR
FLIGHT MANAGEMENT COMPUTER
FLIGHT MANAGEMENT SYSTEM
AUTOFLIGHT STATUS ANNUNCIATOR
FORCE LIMITER CLUTCH ENGAGE
FREQUENCY
FEET PER SECOND PER SECOND
2
MAINTENANCE TRAINING MANUAL
SORTED BITE ABBREV
FT
FUNCT
FWD
G/S
GND
GND FUNCT
GRAD
GS
H/O
HDG
HDGSL/WL
HIS
HYD
HYD FLT CONT SW
IFSAU
ILS
ILS DEV W/R
IN-LBS
INHIB
INPT
INST
INV
IRS SYS DSPL SW
IRU
KILO
KT
LAT ACCEL
LBS
LCL
LIM
LL
LNAV
LNV
LOC ANT RLY
LOC TUN
LONG ACCEL
LRRA
LRU
LVDT
LWD
MTRM
MACH/AS
MAG
MAINT MAN
MAINT MANUAL
MAN
FEET
FUNCTIONAL
FORWARD
GLIDE SLOPE
GROUND
GROUND FUNCTIONAL
GRADIENT
GLIDE SLOPE
HARD OVER
HEADING
HEADING SELECT/WINGS LEVEL
HORIZONTAL SITUATION INDICATOR
HYDRAULIC
HYDRAULIC FLIGHT CONTROL SWITCH
INTEGRATED FLIGHT SYSTEMS ACCESSORY UNIT
INSTRUMENT LANDING SYSTEM
ILS DEVIATION WARNING RESET
INCH-POUNDS
INHIBIT
INPUT
INSTRUMENT
INVALID
INERTIAL REFERENCE SYSTEM DISPLAY SWITCH
INERTIAL REFERENCE UNIT
KILOTON
KNOT
LATERAL ACCELERATION
POUNDS
LOCAL
LIMIT, LIMITER
LOWER LIMIT
LATERAL NAVIGATION
LATERAL NAVIGATION
LOCALIZER ANTENNA RELAY
LOCILIZER TUNED
LONGITUDINAL ACCELERATION
LOW RANGE RADIO ALTIMETER
LINE REPLACEABLE UNIT
LINEAR VARIABLE DIFFERENTIAL TRANSFORMER
LEFT WING DOWN
MACH TRIM
MACH AIRSPEED INDICATOR
MAGNETIC
MAINTENANCE MANUAL
MAINTENANCE MANUAL
MANUAL
3
MAINTENANCE TRAINING MANUAL
SORTED BITE ABBREV
MCP
MCP C/O SW
MIN
MSU
MT
MT ACT
MTA
MTA BRK
MTS
MTS SEL STA
N1 POT
NAV
NOM
NORM ACCEL
NSS
OP
OV/RD
P/B
P INTG RST
PADDLE WRPH
PANL
PAR
PBOV
PCU
PCWS
PGM
PITOT/STAT PROBES
PLCRD
POS
POT
PRESS
PRI
PSIG
PUSHBTN
R INTG RST
R LOC ANT
RADIO ALT
RBOV
RCVR
RCWS
RLY
RST
SEGS
SEL
SEN
SENS
MODE CONTROL PANEL
MODE CONTROL PANEL CHANGEOVER SWITCH
MINIMUM
MODE SELECTOR UNIT
MACH TRIM
MACH TRIM ACTUATOR
MACH TRIM ACTUATOR
MACH TRIM ACTUATOR BRAKE
MACH TRIM SYSTEM
MACH TRIM SYSTEM SELECT STATUS
N1 POTENTIOMETER
NAVIGATION
NOMINAL
NORMAL ACCELERATION
NEUTRAL SHIFT SENSOR
OPTION
OVERRIDE
PUSH BUTTON
PITCH INTEGRATOR RESET
PADDLE WRAP HIGH
PANEL
PARITY
PITCH BAR OUT OF VIEW
POWER CONTROL UNIT
PITCH CONTROL WHEEL STEERING
PROGRAM
PITOT STATIC PROBES
PLACARD
POSITION
POTENTIOMETER
PRESSURIZE
PRIMARY
POUNDS PER SQUARE INCH GAUGE
PUSHBUTTON
ROLL INTEGRATOR RESET
REAR LOCALIZER ANTENNA
RADIO ALTIMETER
ROLL BAR OUT OF VIEW
RECEIVER
ROLL CONTROL WHEEL STEERING
RELAY
RESET
SEGMENTS
SELECT
SENSOR
SENSOR
4
MAINTENANCE TRAINING MANUAL
SORTED BITE ABBREV
SOP
SPD
SPD TRM
SPD/ALT INTV
SPOIL
STAB
STB TRM CUT
STD
STK
STR
SUP
SUPFLG
SURF
SW
SYS
T DWN
T UP
T.E.
TE
TND
TOGA SW
TRIM WARN ANN
TRM CLUCH W
TRM/FLP LK
TRM/FLP TLK
TRUE/MAG
TRUE/MAG HDG SW
TST
UL
UP/LT
V/S
VAC
VAL
VERT
VHF/NAV CONT PANL
VMO/MMO
VNAV
VOR
WHL LIM IN
WRN
WRP
X-CHAN
XDCR
XFR
XMTR
STANDARD OPTION PIN
SPEED
SPEED TRIM
SPEED/ALTITUDE INTERVENTION
SPOILER
STABILIZER
STABILIZER TRIM CUTOUT
STANDARD
STICK
STRETCHED
SUPERFLAG
SUPERFLAG
SURFACE
SWITCH
SYSTEM
TRIM DOWN
TRIM UP
TRAILING EDGE
TRAILING EDGE
TUNED
TAKEOFF GO AROUND SWITCH
STAB OUT OF TRIM ANNUNCIATOR
STAB TRIM CLUTCH WRAP
TRIM/FLAP INTERLOCK
TRIM/FLAP INTERLOCK
TRUE/MAGNETIC
TRUE/MAGNETIC HEADING SWITCH
TEST
UPPER LIMIT
UP/LEFT
VERTICAL SPEED
VOLTS ALTERNATING CURRENT
VALID
VERTICAL
VHF NAVIGATION CONTROL PANEL
MAXIMUM OPERATING SPEED/MAXIMUM OPERATING MACH
VERTICAL NAVIGATION
VHF OMNIRANGE
WHEEL LIMIT IN
WARN
WRAP
CROSS CHANNEL
TRANSDUCER
TRANSFER
TRANSMITTER
5
EFIS
1
MAINTENANCE TRAINING MANUAL
INTRODUCTION
ELECTRONIC FLIGHT INSTRUMENT SYSTEM
Purpose
The Electronic Flight Instrument System (EFIS) provides displays for most of the airplane
navigational systems.
The EFIS provides color displays of: pitch and roll, navigational maps, weather, radio
altitude and decision height and autopilot and flight path information. It also provides
displays of: airspeed, ADF/VOR bearings, ILS data and stall warning information.
System Description
This system includes the Electronic Attitude Director Indicators (EADI's), Electronic
Horizontal Situation Indicators (EHSI's), EFIS Symbol Generators (SG's), EFIS Control
Panels, and EFI Transfer Switch and associated relays.
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MAINTENANCE TRAINING MANUAL
ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)
3
MAINTENANCE TRAINING MANUAL
COMPONENT LOCATIONS
EFIS COMPONENT LOCATIONS
General Component Locations
EADIs and EHSIs
The EADIs and EHSIs are located directly in front of the captain and first officer on
the captain's and first officer's instrument panels (P1 and P3).
EFIS Control Panels
Separate control panels for the captain and first officer are located on the left and
right side of the aft electronic panel (P8).
EFIS Symbol Generators
The left and right EFIS symbol generators are located in the electronic equipment (E/E)
compartment on rack E2-3.
EFI Transfer Switch
The EFI transfer switch is located on the pilots' overhead panel (P5-Fwd).
EFIS Transfer Relays
The EFIS transfer relays are located in the E/E compartment on rack E1-2.
Remote Light Sensors
Two light sensors are located on the right and left sides of glare shield panel, P7.
4
MAINTENANCE TRAINING MANUAL
EFIS COMPONENT LOCATIONS
5
MAINTENANCE TRAINING MANUAL
NOTES
6
MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
EFIS CONTROL PANEL
Purpose
The EFIS Control Panel controls display modes and ranges, allows selection of Decision
Height (DH), and provides the weather radar system ON/OFF control.
Features
EADI Controls
BRT — controls brightness level of EADI display.
DH REF — this liquid crystal display shows the selected decision height.
Decision height set knob — this 24-detent, continuous-turn control knob selects
the decision height.
RST — manually resets DH-Alert on the EADI.
EHSI Controls
RANGE — selects the range for the navigation, Traffic Alert and Collision
Avoidance (TCAS), and weather radar data display on the EHSI for the flight
management computer and for the weather radar receiver transmitter. The TFC
button on the range control allows TCAS traffic to be displayed on the EHSI when
activated.
Mode select switch — selects mode of data to be displayed on the EHSI.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
EHSI Controls (Cont)
BRT — these are two concentric knobs. The outer controls the overall brightness of the
EHSI display; the inner controls the relative brightness of the weather radar display.
WXR — Weather Radar ON/OFF Switch. When this switch is pressed by the captain
or first officer, it turns on the weather radar receiver transmitter unit and enables
the display of weather radar information on their associated EHSI. The switch is a
push-on/push-off type and illuminates when pushed on.
MAP Mode Display Selector Switches — In the MAP, CTR MAP, or PLAN modes, these
switches cause the display of the symbols listed below. Any or all MAP display
switches may be actuated at the same time. The switches are push-on/push-off
and illuminate when actuated. There is a white band around the rim of each
pushbutton cap that is visible only in the OFF position.
VOR/ADF — VOR and ADF bearing data.
MAP modes.
Turns on VOR and ADF vectors in MAP or CTR
NAV AID — VOR, VORTAC, etc. All the navigation aids are displayed
in the three lowest range selections. Only the high altitude navigation aids are
displayed in the upper three range selections.
ARPT — airports.
RTE DATA — altitude restrictions and estimated time-of-arrival.
WPT — waypoint not in the selected flight plan. Waypoints are displayed only in
the lowest three range selections.
Power
The EFIS Control Panel is powered by 115 volt ac, 400 Hz.
8
MAINTENANCE TRAINING MANUAL
EFIS CONTROL PANEL
9
MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
EFIS SYMBOL GENERATOR
Purpose
The EFIS symbol generator processes data from the EFIS control panel and navigation and
guidance systems to provide video signals to the EADI and EHSI displays.
Physical Description
The symbol generator weighs 21 pounds (9.5Kg) and is packaged in a size 6 MCU case.
Features
Front Panel
The TEST switch initiates a self-test for checking the Symbol Generator, Display Units,
and Control Panel. The RESET switch is inoperative in the Boeing 737.
Operation
Input data is received on many ARINC 429 data buses, as well as ARINC 453 weather radar
data buses and discrete inputs. Program pins allow for changes to certain system
displays.
The main processing circuitry includes a central processing unit, memory units, raster
and stroke generators, and input/output buffer circuits. Six different colors can be
generated at two different intensity levels.
BITE
BITE circuitry detects faults and identifies faulty line replaceable units. Failure data is
stored in nonvolatile memory.
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MAINTENANCE TRAINING MANUAL
EFIS SYMBOL GENERATOR
11
MAINTENANCE TRAINING MANUAL
NOTES
12
MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
EFIS — EADI AND EHSI
Purpose
The EADI and EHSI are the display units for EFIS.
Physical Description
The EADI weighs 23 pounds (10.8 kg) and is 6 inches wide x 5.5 inches high x 14 inches deep
with a display area 4.7 inches wide x 4.2 inches high. The EHSI weighs 25 pounds (11.3 Kg)
and is 6 inches wide x 7 inches high x 14 inches deep and display area 4.7 inches wide x 5.7
inches high. Both units require cooling air from the equipment cooling system through inlet
and exhaust holes on the rear of their case. Wire mesh screens cover these holes to prevent
lint and foreign debris from entering the display units. These screens should be checked
regularly for buildup of foreign debris.
Operation
General
Each display unit uses a cathode ray tube (CRT) for the display. The CRTs use 115 volt
ac, 400 Hz, power. An EFIS Symbol Generator (SG) provides analog deflection and
digital video signals to the CRT.
There are three electron guns, one each for red, green, and blue. By controlling the
three guns, the symbol generators can produce the colors red, green, and blue, as well
as combinations resulting in the colors yellow, cyan (light blue), magenta (pink), and
white. If one or more electron guns fail, the display will revert to a monochromatic
display.
The EADI's inclinometer serves as a slip indicator using a black ball against a white
background.
Scanning Methods
Two types of scanning are used: raster and stroke. The vertical raster is used on the
EADI to paint the attitude ball and on the EHSI for the weather radar display. All
other symbols use the point-to-point stroke type of scan. The displays are made
flicker-free by refreshing the stroke-written symbols at the rate of 80 Hz and the
raster at the rate of 40 Hz. The data is processed serially, first the EADI, then the EHSI,
alternately. An overtemperature condition in either display is indicated by the removal
of the raster video at the first overtemperature threshold and a blank display at the
second overtemperature threshold.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
Maintenance Practices
The display units are held in place in mounting racks by hinged handle assemblies and
fastening screws. As the fastening screws are loosened or tightened for removal or
installation, push the handle firmly against the front of the unit. If not pushed in,
threads holding the screws will possibly be stripped and the display unit will be damaged.
The local light sensor on the bottom edge of each display unit automatically maintains the
brightness of the display as ambient light conditions change within the flight compartment
area.
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MAINTENANCE TRAINING MANUAL
EFIS - EADI AND EHSI
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COMPONENT FUNCTIONAL DESCRIPTION
EFIS REMOTE LIGHT SENSORS
Purpose
Two light sensors, mounted on the glareshield, are used to change the display intensity of
the EADI and EHSI displays.
Physical Description
Each sensor is approximately two inches wide, one inch long, and one inch high and has a
field of view of 45° left and right and 15° up and down.
Operation
The sensors face forward so that their photo diodes measure ambient light coming through the
windshield. The output is an analog signal proportional to external ambient light. Each
unit is used in the control circuit for the onside EADI-EHSI pair. Power to each sensor is
+24 volt dc and -5 volt dc from its associated EHSI.
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MAINTENANCE TRAINING MANUAL
EFIS REMOTE LIGHT SENSOR
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
EFIS — TRANSFER SWITCH
Purpose
Normally, the left symbols generator is the source for the left displays and the right
symbol generator is the source for the right displays. If a symbol generator fails (or the
flight crew decide for another reason), the EFIS transfer switch can be used to allow one
symbol generator to be the source for all the displays.
When the switch is in the "BOTH ON 1" position, both EADIs and EHSIs are driven by the left
symbol generator. In the "BOTH ON 2" position, they are all driven by the right symbol
generator.
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MAINTENANCE TRAINING MANUAL
EFI TRANSFER SWITCH
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS — TYPICAL EADI DISPLAY
Operation
The EADI has one basic display mode. The EADI displays airplane attitude, flight director
commands, Mach, airspeed, and ILS and radio altimeter data. Across the top of the EADI,
autothrottle and autopilot annunciations are provided for both "armed" and "engaged"
conditions.
An inclinometer is installed at the bottom of the EADI to provide turn-and-bank
information.
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MAINTENANCE TRAINING MANUAL
EFIS – TYPICAL EADI DISPLAY
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MAINTENANCE TRAINING MANUAL
On the following sheets, an illustration is provided of each possible EADI symbol along with
its name, display location, color, description and data source.
The EADI options selected are:
AIRPLANE SYMBOL ...........................................
SPEED TAPE ................................................
SKY/GROUND ................................................
PITCH SCALE ...............................................
FLIGHT DIRECTOR ...........................................
HEIGHT ALERT ..............................................
GROUND SPEED ..............................................
MACH......................................................
TRUE AIRSPEED .............................................
SINGLE CHANNEL ANNUNCIATION ...............................
WINDSHEAR WARNING .........................................
PITCH/ROLL COMPARATOR .....................................
FLASHING PITCH/ROLL COMPARATOR ............................
RADIO ALTITUDE ............................................
RISING RUNWAY .............................................
PITCH LIMIT ...............................................
RESOLUTION ADVISORY .......................................
AIRSPEED TAPE OPTIONS
LOW SPEEDS BOTTOM OF TAPE ............................
CAS INDICATOR ........................................
AIRSPEED TREND VECTOR ................................
MIN OPERATING SPEED ..................................
EFIS – NORMAL EADI DISPLAY SYMBOLS
22
SPLIT AXIS
YES
SPEED TAPE
SPEED TAPE
COMMAND BAR
NO
YES
YES
NO
YES
YES
YES
NO
DIGITAL + ANALOG DIAL
YES
YES
YES
YES
ROLLING DIGIT CURSOR
YES
YES
MAINTENANCE TRAINING MANUAL
EFIS - NORMAL EADI DISPLAY SYMBOLS
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MAINTENANCE TRAINING MANUAL
EFIS - NORMAL EADI DISPLAY SYMBOLS
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EFIS - NORMAL EADI DISPLAY SYMBOL
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EFIS - NORMAL EADI DISPLAY SYMBOL (SPEED TAPE)
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS - EADI FAULT DISPLAYS
Operation
Whenever specific input data to a symbol generator is identified as "invalid" data, the
associated symbol, or parameter value, is blank. In many cases, yellow fault flags are
displayed as well. The location, orientation, and letter characters associated with each of
these flags are shown.
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MAINTENANCE TRAINING MANUAL
EADI - FAULT DISPLAYS
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS - EADI FAULT AND NCD DISPLAYS
Operation
On the following sheets, the various symbol or parameter display changes are described for
both "invalid" and "no-computed-data" (NCD) conditions.
For NCD conditions, the associated symbols, or parameter values, are removed from the EADI
display. In some cases, parameter values are replaced by a series of dashes which are
located in the same position as for normal data. For the locations, refer to the previously
shown illustration of a typical EADI display.
For conditions where input data to the symbol generators is invalid, the associated symbols,
or parameter values, are removed from the display. In most cases, an appropriate yellow
fault flag is also displayed.
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EFIS - EADI INVALID DATA AND NCD DISPLAY (SHEET 1)
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EADI - INVALID DATA AND NCD DISPLAY (SHEET 2)
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NOTES
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OPERATION
EFIS - TYPICAL EHSI DISPLAYS
Operation
VOR/ILS Mode Displays
The VOR and ILS modes are used with manually selected stations. The VOR mode is used
enroute with VOR and DME station inputs. The ILS mode is used for landing and uses ILS
and DME station inputs. Appropriate deviations are displayed in both modes.
Heading displayed is magnetic between 73° north and 60° south latitude. At higher
latitudes, the heading is automatically referenced to true North.
The drift angle can be determined by computing the difference between the heading and
track lines.
For the VOR and ILS systems, the EHSI has two different types of displays (FULL and EXP
VOR/ILS) which are selectable on the EFIS Control Panel.
The "FULL VOR/ILS" display includes a 360-degree compass rose with the airplane symbol
in the center.
The "EXP VOR/ILS" display includes a 70-degree compass arc with the airplane symbol at
the bottom center. This display is used to more easily read the airplane's heading and
track values. Also, the weather radar data can be displayed on the expanded display.
Whether the display presents VOR or ILS data depends upon the VOR or ILS frequency
selection on the VHF NAV Control Panel. The frequency is displayed on the bottom of
the EHSI.
NAV Mode Displays
The NAV modes display selected FMC data. The display shows the aircraft position
relative to the next waypoint and the vertical and horizontal flight path.
Like the VOR/ILS modes, the NAV modes have selectable 360-degree and 70-degree
compass displays.
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EFIS - TYPICAL EHSI DISPLAYS
MAINTENANCE TRAINING MANUAL
OPERATION
EFIS-TYPICAL EHSI CENTER MAP, MAP, AND PLAN DISPLAYS
Operation
The EHSI can display selected FMC data in three different selectable formats: CTR MAP, MAP,
and PLAN. The format displayed depends upon the mode selected on the EFIS Control Panel.
PLAN Mode
This mode is primarily used prior to flight to set up the flight plan. The upper part
of the display describes the dynamic conditions of the airplane: track, selected and
actual heading, and distance and estimated time of arrival to the next waypoint. The
lower part of the display is background data displaying the flight plan. The content
of this display is determined by switch selections on the FMC Control Display Unit
(CDU) and the EFIS control panel.
During PLAN mode, the display is oriented "north-up", and there is no weather radar
display.
MAP and CTR MAP Modes
These modes display the airplane's position with respect to the flight plan. The
entire display is dynamic, so it translates and rotates as the airplane moves. The
airplane is represented on the map by the triangle at the center of the display.
Waypoint (WPT) and NAVAID symbols remain upright, but airport (ARPT), holding pattern,
and procedural turn symbols are rotated to maintain proper orientation.
These display modes are generally used during flight to monitor airplane position along
the selected flight path stored in memory. The Center Map Mode display includes a 360degree compass rose. The Map Mode display includes only a 70-degree compass arc which
is used to more easily monitor the airplanes heading, track, and progress along the
selected route. Weather radar data can be shown in both MAP modes.
Except for the differences noted, the remainder of the navigational data shown is the
same on both types of displays. A detailed description of all displayed symbols,
parameters, and annunciations is provided later.
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MAINTENANCE TRAINING MANUAL
EFIS - TYPICAL MAP, CENTER MAP, AND PLAN DISPLAYS
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MAINTENANCE TRAINING MANUAL
On the following sheets, an illustration is provided of each possible EHSI symbol along with
its name, display location, color, description and data source.
The EHSI options selected are:
MODES AVAILABLE ........................... EXPANDED AND FULL VOR/ILS AND NAV,
MAP, CENTER MAP AND PLAN
CENTER MAP FORMAT ......................... FULL ROSE
HEADING/TRACK INDICATIONS:
NAV MODES ........................... TRACK UP
VOR/ILS MODES ....................... HEADING UP
MAP MODES ........................... TRACK UP
PLAN MODE ........................... TRACK UP
SYMBOLOGY ................................. BASIC
RADAR RANGE INDICATION .................... RANGE MARKS
ADF POINTERS IN MAP MODES ................. YES
VOR/DME AUTOTUNE .......................... YES
FMC/IRU POSITION DIFFERENCE ............... YES, DISCRETE FROM FMC AT 4NM
LEFT ANCDU ................................ NO
RIGHT ANCDU ............................... NO
FMC....................................... DUAL
TRUE AIRSPEED ............................. NO
GROUND SPEED .............................. NO
RADAR TURBULENCE .......................... MAGENTA
MAP MODE ANNUNCIATION ..................... NO
SELECTED COURSE FROM VOR
STATION DISPLAYED IN MAP ................ YES
TCAS STATUS, TRAFFIC ONLY, TRAFFIC,
OFF SCALE MESSAGE, TRAFFIC SYMBOLS,
TRAFFIC DATA, TRAFFIC ARROW,
NO BEARING TRAFFIC ...................... YES
TCAS RANGE RING ........................... NO
EFIS - NORMAL EHSI DISPLAYS SYMBOLS
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MAINTENANCE TRAINING MANUAL
EFIS - NORMAL EHSI DISPLAYS SYMBOLS
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EFIS - NORMAL EHSI DISPLAYS SYMBOLS
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EFIS - NORMAL EHSI DISPLAYS SYMBOLS
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MAINTENANCE TRAINING MANUAL
EFIS - NORMAL EHSI DISPLAYS SYMBOLS
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EFIS - NORMAL EHSI DISPLAYS SYMBOLS
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS - EHSI INVALID DATA ANNUNCIATIONS
Operation
Whenever specific input data to a symbol generator is identified as "invalid" data, the
associated symbol, or parameter value, is blanked. In many cases, yellow fault flags are
displayed as well. The location, orientation, and letter characters associated with each of
these flags are shown.
The yellow fault flags will be displayed under the following conditions:
FLAG
FAILED COMPONENT
HDG
TRK
XTK
VTK
MAP
VOR
LOC
TCAS
FAIL
IRU
FMC AND IRU
FMC
FMC
FMC
DAA
DAA
TCAS COMPUTER
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MAINTENANCE TRAINING MANUAL
EFIS - EHSI INVALID DATA ANNUNCIATIONS
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS - EHSI FAULT AND NCD DISPLAYS
Operation
On the following sheets, the various symbol or parameter display changes are described for
both "invalid" and "no-computed-data" (NCD) conditions.
For NCD conditions, the associated symbols, or parameter values, are generally removed from
the EHSI display. In some cases, parameter values are replaced by a series of dashes which
are located in the same position as for normal data. For these locations, refer to the
previously shown illustrations of typical EHSI displays.
For conditions where input data to the symbol generators is invalid, the associated symbols,
or parameter values, are removed from the display. In most cases, an appropriate yellow
invalid data flag is also displayed.
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MAINTENANCE TRAINING MANUAL
EFIS - EHSI FAULT AND NCD DISPLAYS (SHEET 1)
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EFIS - EHSI INVALID DATA AND NCD DISPLAYS (SHEET 2)
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EFIS - EHSI INVALID DATA AND NCD DISPLAYS (SHEET 3)
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS BLOCK DIAGRAM
Operation
General
The block diagram shows EFIS inputs and its interface with other systems.
EFIS Configuration
EFIS consists of two symbol generators, two control panels, two electronic horizontal
situation indicators, two electronic attitude director indicators, two remote light sensors,
and an EFI transfer switch.
Power
Each EFIS component receives 115 volt ac power from an individual circuit breaker.
Switching power is 28 volt dc.
Control
EFIS control is achieved at two EFIS control panels, one connected to the left symbol
generator, the other connected to the right symbol generator.
Inputs
Sensor inputs are of the following types: ARINC 453 digital data buses, ARINC 429 digital
data buses, dc-analog, synchro, and discretes. The inputs are grouped on the graphic as
switched and unswitched inputs. A group of unswitched inputs is connected to both symbol
generators. Two other groups of unswitched inputs are connected to only the left or the
right symbol generator.
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MAINTENANCE TRAINING MANUAL
OPERATION
Video
Video outputs are sent from the left symbol generator to the left display units, and from the
right symbol generator to the right display units. Two video cross-ties between symbol
generators allow selecting either the left or the right symbol generator as the video data
source. The EFI transfer switch and relays select the symbol generator to be used.
Digital Output
An ARINC 429 output bus from each symbol generator is connected to the other symbol generator,
the flight management computer, and the digital flight data recorder system. This bus carries
attitude (for comparison), radio altitude, BITE, and discrete status words.
Display Unit Brightness
Display unit brightness is controlled manually from the EFIS control panels, and automatically
by remote and ambient light level sensors.
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EFIS BLOCK DIAGRAM
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS SYMBOL GENERATOR INPUTS, EFIS CONTROL PANEL, WEATHER RADAR, AND FLIGHT MANAGEMENT COMPUTER
Operation
The following graphic illustrates the EFIS system input interfaces with the weather radar
system and the flight management computer system.
Weather Radar System
The weather radar systems can be turned on from either EFIS control panel. The on/off
discrete is sent through the weather radar control panel, which provides an interlock (if
the weather control panel is not installed, or is not powered the weather radar system will
not turn on). The weather radar on signal also enables the EFIS symbol generator display of
weather radar information. The EFIS control panel also provides range to the weather radar
system. The captain and first officer can select different ranges. The weather radar
receiver/transmitter provides weather information, weather mode, antenna tilt, and radar
gain to the EFIS symbol generator.
Flight Management Computer System
The EFIS control panel supplies range, mode, and MAP switch selection to the flight
management computer system. The flight management computer system supplies route and map
navigation data on the EHSI and target speeds and ground speed on the EADI.
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MAINTENANCE TRAINING MANUAL
EFIS - EFIS CONTROL PANEL/WEATHER RADAR/
FLIGHT MANAGEMENT COMPUTER INPUTS
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS - RADIO ALTIMETER INPUTS
Operation
The radio altimeter system provides the symbol generator with radio altitude. The symbol
generators display the data on the EADI's whenever the 28 volt dc radio altitude valid
signal is also being received.
The EFIS control panel provides a decision height value to the EFIS symbol generators.
Decision Height (DH) is displayed along with the actual radio altitude. When the airplane
descends below the decision height, a DH alert is annunciated. The DH alert will continue
to be displayed until reset. DH alert is a function of the EFIS symbol generators.
The rising runway option is a green representation of an airport runway that is present when
an active, valid localizer frequency is selected on the VHF NAV control panel. The runway
is parked at the bottom of the EADI until the airplane descends below 200 feet radio
altitude. As the airplane descends below 200 feet, the runway rises until landing.
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MAINTENANCE TRAINING MANUAL
EFIS SYMBOL GENERATOR – RADIO ALTIMETER INPUTS
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS SYMBOL GENERATOR - VHF NAV/DME INPUTS
Operation
DME and VOR/ILS data are sent to the Digital to Analog Adapters (DAA) in an analog format.
In the DAA's, the data is changed into an ARINC 429 data bus format and sent to the EFIS
symbol generators. The glide slope data is sent in an analog format to the symbol
generators through the VHF NAV transfer relay. The data is processed by the symbol
generators and displayed when the VOR/ILS mode is selected on the EFIS control panel.
Normally the captain's EADI and EHSI display localizer information from the number one DAA
and glide slope information from the number one VHF NAV receiver, while the first officer's
EADI and EHSI display data from the number two DAA and VHF NAV receiver. VOR data, DME
distance, and the VOR/ILS frequency are similarly displayed on the captain's and first
officer's EHSI. If the VHF NAV transfer switch is moved to the BOTH on 1 or BOTH on 2
position, the VOR, the ILS, and DME displays will be from the selected system on all four
EFIS displays.
The VOR vectors are displayed on the EHSI's in the MAP and CTR MAP modes if the VOR/ADF
select switch is pressed on the EFIS control panel.
The digital flight control system mode control panel sends the VOR selected course or
selected runway heading to the EFIS symbol generators on a digital data bus. (This data bus
is not shown on the graphic.) The symbol generator combines VOR-selected course and VOR
bearing and computes VOR deviation and TO/FROM information.
The EFIS EADI display circuits use the VOR/ILS frequency from the DAA to determine whether
to display ILS information. The EFIS EHSI circuits use the ILS 28 volts DC signal from the
EFIS control panel to determine whether to display VOR or ILS information. Because of this
difference, if the DAA fails, the EADI will always display a LOC flag, but the EHSI will
display the appropriate flag (VOR or LOC) depending on the frequency selected on the VHF NAV
control panel.
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MAINTENANCE TRAINING MANUAL
OPERATION
Signal Types
DAA Output Signals
The DAA-1 and DAA-2 VOR, LOC and DME data is sent to both EFIS symbol generators on ARINC
429 digital data buses.
DME Output Signals
The distance data to the DAAs is on ARINC 568 digital data buses.
The DME valid signal is 28 volts dc.
VHF Control Panel Signals
The VOR/ILS frequency data is sent to the DAAs using 2-out-of-5 discretes.
VHF NAV Receiver Signals
Outputs from the VHF NAV receiver are as follows:
VOR bearing in a synchro format to the DAA
LOC deviation in an analog format to the DAA
Glide slope deviation in an analog format to the VHF NAV transfer relays
VOR/LOC superflag, 28 volts dc, to the DAA
Glide slope superflag, 28 volts dc, to the VHF NAV transfer relays
ILS 28 volts dc to the EFIS control panels via the VHF NAV transfer relays and
to the DAA
The EFIS symbol generators display left or right VOR, LOC, and DME data depending
on the state of the input source select discretes (open = left and ground =
right).
EFIS Control Panel
The mode selection data is sent to the EFIS symbol generator on an ARINC 429
digital data bus.
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MAINTENANCE TRAINING MANUAL
EFIS SYMBOL GENERATOR - VHF NAV/DME INPUTS
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OPERATION
EFIS SYMBOL GENERATOR – IRS INPUTS
Operation
IRS data is sent on Bus No. 1 and Bus No. 3 from IRU-1 and IRU-2 to both symbol generators
and to other user systems. Both buses are ARINC 429 digital data buses.
The EADI uses pitch and roll data for its pitch angle and roll angle readouts and also for
positioning the horizon line. The EHSIs use IRU heading data in all modes and in case of an
FMC failure, they also use the IRU generated track angle, wind, ground speed and drift angle
parameters.
IRU Switching
As soon as airplane power is applied, the symbol generator switch S1 will actuate
causing the left IRU Data Bus No. 1 to provide data to the left symbol generator, and
the right IRU Data Bus No. 1 to provide data to the right symbol generator.
If an IRU fails, the IRS transfer relay may be energized to the "BOTH ON 1" or "BOTH ON
2" position to provide IRS data to both symbol generators from one IRU.
Present position data from both IRU inputs is used in both symbol generators to
establish FMC/IRU position difference for both IRUs.
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EFIS SYMBOL GENERATOR – IRS INPUTS
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OPERATION
EFIS SYMBOL GENERATOR – AIR DATA COMPUTER INPUTS
Operation
The air data computers provide air data to the EFIS symbol generators on an ARINC 429
digital data bus. The symbol generators process the input data from their onside ADC. The
symbol generators use these inputs to control the speed tape and to display MACH on the
EADI.
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MAINTENANCE TRAINING MANUAL
EFIS SYMBOL GENERATOR – AIR DATA COMPUTER INPUTS
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MAINTENANCE TRAINING MANUAL
OPERATION
EFIS SYMBOL GENERATOR – FCC/MODE CONTROL PANEL INPUTS
Operation
General
The autoflight director system input data to the left EFIS symbol generator normally
comes from flight Control Computer (FCC)-A on the FCC-L ARINC 429 digital data bus. The
parameters and discretes listed on the graphic are used by the EFIS symbol generator.
In a similar manner, data to the right EFIS symbol generator will come from FCC-B on the
FCC-R digital data bus.
Selected Heading and Course Parameters
If the selected course No. 1 data word received by the left EFIS symbol generator on
the FCC-L bus is invalid, the symbol generator will switch over to the MCP-1 digital
data bus to receive this parameter. In this case, the course data word is received
directly from the DFCS Mode Control Panel (MCP), rather than indirectly through FCC-A.
If the selected heading data word received by the symbol generator on the FCC-L bus is
invalid, the symbol generator will switch over to FCC-R bus, if that input is invalid
then the symbol generator will switch over to the MCP-1 data bus. The right EFIS symbol
generator will receive SEL HDG and SEL CRS No. 2 from data buses FCC-B or MCP-2 in a
similar manner.
Flight Director Parameters
The left EFIS symbol generator's source for the Flight Director ON/OFF discrete and the
Flight Director Pitch/Roll bar drive signals is FCC-A. If this source is switched off,
or becomes invalid, the flight director bars are removed.
Selected Target Speed
Both EFIS symbol generators use the selected target speed value received on FCC-L bus
or FCC-R bus, depending upon which FCC is the "master".
DFCS Modes and Status
Both EFIS symbol generators utilize the autoflight mode and status data from the
"master" FCC received across buses FCC-L or FCC-R.
Localizer/Glide Slope Deviation Warnings
The EFIS symbol generators will display localizer and/or glide slope deviation warnings
based on discrete signals sent by the onside FCC on a data bus.
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MAINTENANCE TRAINING MANUAL
EFIS – SYMBOL GENERATOR – FLIGHT CONTROL COMPUTER/
MODE CONTROL PANEL INPUTS
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OPERATION
EFIS SYMBOL GENERATOR – AUTOTHROTTLE COMPUTER INPUTS
Operation
The autothrottle computer sends parameters on an ARINC 429 digital data bus A/T-1 to the
EFIS symbol generators, the flight control computers, and the flight management computer
system (FMCS).
In addition an analog discrete, "autothrottle computer installed", is sent to both EFIS
symbol generators to indicate that the autothrottle computer is installed. Also, a
"throttle hold" analog discrete is received during the takeoff phase when the autothrottle
computer enters the throttle-hold mode. When this occurs, the "THR HLD" annunciation appears
on the EADIs.
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MAINTENANCE TRAINING MANUAL
EFIS SYMBOL GENERATOR – AUTOTHROTTLE COMPUTER INPUTS
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OPERATION
EFIS SYMBOL GENERATOR – AUTOMATIC DIRECTION FINDER INPUTS
Operation
Power
ADF 1 and ADF 2 receive power via the ADF control panel. The "ADF 1 ON" and "ADF 2 ON"
discretes are sent to both EFIS symbol generators.
ADF Bearing Signals
The ADF receiver sends ADF bearing data, along with a 26 volt ac reference signal, to
both EFIS symbol generators in an analog format.
The EFIS displays the ADF bearing data on the EHSI utilizing pointer symbols in the
VOR/ILS modes, NAV modes, Map modes and PLAN mode. ADF pointers are removed and ADF
bearing vectors are displayed in the Map modes when the VOR/ADF key on the EFIS control
panel is pushed.
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EFIS Symbol Generator – ADF Inputs
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OPERATION
EFIS SYMBOL GENERATOR – STALL WARNING COMPUTER, GROUND PROXIMITY WARNING, AND TRAFFIC ALERT AND
COLLISION AVOIDANCE INPUTS
Operation
Stall warning computer 1 provides the following data values to the left EFIS symbol
generator on an ARINC 429 digital data bus: 1) minimum operating speed; 2) stick shaker
speed; 3) engine out speed; 4) minimum flap retraction speed; 5) high speed buffet; 6)
maximum operating speed; and 7) pitch limit angle. In a similar manner, stall warning
computer 2 provides the same data values to the right EFIS symbol generator.
The ground proximity warning computer sends a windshear alert discrete to both EFIS symbol
generators when windshear conditions are sensed.
The TCAS computer send resolution advisory signals to both symbol generators to be displayed
on the EADIs. Likewise traffic advisories and TCAS status are sent to both symbol
generators for display on the EHSIs.
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EFIS SYMBOL GENERATOR – STALL WARNING COMPUTER, GROUND PROXIMITY WARNING SYSTEM AND
TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM
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OPERATION
EFIS – SYMBOL GENERATOR SELECTION
Control Sequence
During normal operation, the left EFIS symbol generator's output drive select relay sends
processed EHSI and EADI data from the EFIS symbol generator to the captain's EADI and EHSI.
The right EFIS symbol generator works similarly, sending data to the first officer's
instruments.
Backup Operation
If the left EFIS symbol generator fails, the EFI switch could be moved to the "BOTH ON 2"
position. This would energize and latch the EFIS transfer relay-1. When the transfer
relay latches in the new position, power is removed from its electromagnetic coil. In the
"BOTH ON 2" position, a relay contact completes a circuit to energize the left EFIS symbol
generator's output drive select relay. Another contact sends a ground to the left symbol
generator to power down its internal electronics and tell the right symbol generator that
the left symbol generator is powered down. The right symbol generator will not record the
left symbol generator as failed when it receives this discrete. A third contact is used to
ensure that both relays cannot be energized at the same time.
The flight control computers (FCC) use pitch attitude (local) from one IRU and roll attitude
(foreign) from the other IRU. If one of the IRU's is no longer being displayed to the
flight crew because either the IRS transfer relay or the EFIS transfer relay is switched,
the FCC will disengage and begin using the other IRU data.
Two diodes, located in the flight instrument accessory unit, are connected across the EFIS
transfer relays to provide high voltage suppression whenever power is removed from the
relays' electromagnetic actuation coils.
Operation of the switching for the right EFIS symbol generator, utilizing transfer relay-2,
occurs in a similar manner whenever the EFI switch is moved to the "BOTH ON 1" position.
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EFIS – SYMBOL GENERATOR SELECTION
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OPERATION
EFIS SYMBOL GENERATOR'S COMPUTED DISPLAYS
Most of the parameters displayed on the EADI and EHSI use
LRUs. These signals are processed within the EFIS symbol
parameters require several signals in order to compute the
discrete to be displayed. The signals required for these
76
signals that originate from various
generators, and then displayed. Some
appropriate parameter value or
computations are shown on the graphic.
MAINTENANCE TRAINING MANUAL
EFIS SYMBOL GENERATORS' COMPUTED DISPLAYS
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SYSTEM TEST
EFIS BITE INTRODUCTION
Flight Management System (FMS) BITE
FMS BITE permits BITE procedures to be performed on the Flight Management Computer System,
the Digital Flight Control System, the Autothrottle System, the Inertial Reference System,
and the Electronic Flight Instrument System (EFIS). The graphic shows the EFIS interface as
an example; the other subsystems are connected similarly.
Fault Isolation
The FMS BITE system checks the EFIS for inflight faults, ground faults, current faults, as
well as the status of the discretes associated with this system. During these tests, the
CDUs are "locked" together starting with EFIS BITE LEFT, or EFIS BITE RIGHT, or SELF-TEST
selection, and their displays are identical.
Operational Checkout
With the FMC CDUs working through the FMC, BITE requests can be sent to the EFIS symbol
generator being tested. The appropriate BITE responses are sent back through the FMC and
displayed on the CDUs. In addition to BITE checks of the EFIS symbol generators, a selftest of the EFIS can be initiated from either CDU. This self-test causes specific test
patterns to be displayed on the EADIs and EHSIs.
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NOTES
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EFIS BITE INTRODUCTION
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SYSTEM TEST
EFIS - ACCESS TO EFIS BITE INDEX
Test Preparation
With the MAINT BITE INDEX displayed, line select key (LSK) 5L can be pressed to obtain the
EFIS BITE Index. If the both EFIS symbol generator's internal self-test fails, the CDU
indicates that the EFIS BITE is inoperative, and displays a message indicating a need to
check the EFIS SG or its interfaces. The EFIS BITE test index allows the selection of BITE
checks to be performed on the left or right EFIS systems, or a self-test to be performed on
both systems simultaneously. Pressing either EFIS LEFT BITE or EFIS RIGHT BITE will display
the appropriate EFIS BITE TEST FAULT INDEX. The L&R index is shown on the graphic. If the
symbol generator or the interface is invalid and any of the selections on the index is
selected, the display stating "CANNOT ACCESS DATA CHECK SG OR INTERFACE" will be shown.
Pressing LSK 6L (INDEX) returns the display to the MAINT BITE INDEX.
In the example shown, the left EFIS has been selected for BITE checks. When LSK 1L is
pressed, an EFIS-LEFT FAULT INDEX is displayed. This FAULT INDEX allows specific and unique
left EFIS tests to be selected and accomplished.
If LSK 6L on the EFIS LEFT FAULT INDEX page is pressed, the CDU display returns to the EFIS
BITE INDEX page.
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EFIS – ACCESS TO EFIS BITE INDEX
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NOTES
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EFIS BITE INDEX TEST SELECTIONS
SYSTEM TEST
Functional Test
General
After calling up the EFIS LEFT FAULT INDEX, it is possible to display a list of any
"in-flight" or "ground" faults as well as the "current status" and "discrete status".
In-Flight Faults
The EFIS continuously monitors operation of its components and interfaces. For up to
nine consecutive flights, any malfunctions or faults that occur during flight and cause
flight deck effects are stored in a non-volatile memory (NVM). Separate memory
addresses are used for each flight. Readout of this data is possible by selecting the
"IN FLIGHT FAULTS" line select key (2L). In-flight faults are those faults which occur
when ground speed is greater than 20 knots and the air/ground relays indicate in air
state.
Data for each flight is sequentially stored, so on readout, the last flight data in is
the first data out. The last flight is always identified as No. 1, numerically counting
up to the NVM storage limit for the earlier flights.
At the end of a flight, fault data for that flight is labeled as FLT 01. Fault data
for each previous flight is reidentified as the next earlier flight number. Data
relating to the earliest flight is dumped from the memory at the end of the flight
segment. Thus data for the most recent flight will always be labeled FLT 01.
The "in-flight" fault example shown is typical. If more than one page is necessary to
list the faults, the faults are displayed alphabetically, and the pages are identified as
1/2, 2/2, etc. From one to seven pages are available. In this case, the CDU's "NEXT
PAGE" and "PREV PAGE" keys can be used to select the desired pages. If there are no
faults, "NO FAULTS" are displayed across the middle of the display.
Ground Faults
Any faults that first occur while the airplane is on the ground are stored in a
dedicated non-volatile memory. Readout of the data is accomplished by pressing the
"GROUND FAULTS" LSK on the EFIS BITE INDEX. Up to seven pages of alphabetically listed
faults are available. If there are no faults, "NO FAULTS" are displayed across the
middle of the display.
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SYSTEM TEST
Current Status
The present status of all EFIS components and the systems that interface with the EFIS can
be checked by calling up a series of current status pages. This is accomplished by
pressing the "CURRENT STATUS" key (LSK-4L) on the CDU, and then using the CDU's
"NEXT PAGE" key to cycle between the series of status pages. In the example shown,
the status of each "left" or "right" LRU component, which interfaces with the left
EFIS, is shown.
If an LRU fails in-flight and then subsequently clears itself, this intermittent type of
fault is displayed as an intermittent "in-flight" fault. However, on the appropriate
CURRENT STATUS page, the LRU status is reported as "OK".
Discrete Status
The present status of some discretes associated with the EFIS are displayed on a series
of discrete status pages which can be displayed by pressing the "DISCRETE STATUS" Key
(LSK 5L) on the CDU, and using the "NEXT PAGE" Key to cycle between pages.
INDEX Key Selection
If the INDEX Key 6L on any of the "fault" or "status" pages, the display returns to the left
EFIS FAULT INDEX.
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EFIS - SELF-TEST INITIATION AND EADI TEST DISPLAYS
MAINTENANCE TRAINING MANUAL
SYSTEM TEST
EFIS - EHSI TEST DISPLAYS
Operational Checkout
Self-Test Procedures and Displays
Range Switch Operation
As described on the EFIS BITE SELF-TEST page, it is necessary to select all modes
on each control panel in order to completely evaluate the EFIS performance. For
this self-test, it is recommended that the 80 nmi range be selected since the
resultant displays will properly match the test patterns illustrated. If desired,
other ranges can be selected in order to verify operation of the EFIS control panels.
Weather Radar ON/OFF Switch Operational Checks
The EFIS CP's WXR ON/OFF switch should be switched on in order to verify the
display of the magenta, red, yellow, and green sectors which verify the EFIS's
capability for processing and displaying weather radar video data. With the four
weather radar sectors displayed, the control panel's brightness control can be
checked by rotating the inner knob of the EFIS HIS BRT control. Caution should be
exercised when activating the weather radar.
MAP Switch
While the MAP, CTR MAP, or PLAN modes are selected, the various MAP switches on
the EFIS control panel can be actuated to verify that the proper symbol is shown.
During self-test, the RTE DATA switch is inoperative. The VOR/ADF switch causes
the V1, V2, A1, and A2 vectors to be displayed. The NAV AID switch causes the
three VOR, VTC (VOR TACAN), and DME symbols to be displayed. The ARPT switch
causes the single APT (airport) symbol to be displayed. The WPT switch causes a
GRP (ground reference point) symbol to be displayed.
Test Values
All of the data values shown are simulated test values generated within the EFIS
SG and are not dependent upon proper operation of the interfacing systems.
Therefore, the values observed during self-test should coincide with those shown on
the graphic.
TCAS Symbols
If the TFC button is activated on the EFIS control panel, the TCAS symbols shown
on page two will be displayed. TCAS TEST and TRAFFIC will be displayed in all
modes regardless of the position of the TFC button. To properly see these EFISgenerated symbols, it is suggested that a lower range be selected.
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EFIS - EHSI TEST DISPLAYS
MAINTENANCE TRAINING MANUAL
SYSTEM TEST
EFIS - EHSI TEST DISPLAYS
Operational Checkout
Self-Test Procedures and Displays
Range Switch Operation
As described on the EFIS BITE SELF-TEST page, it is necessary to select all modes
on each control panel in order to completely evaluate the EFIS performance. For
this self-test, it is recommended that the 80 nmi range be selected since the
resultant displays will properly match the test patterns illustrated. If desired,
other ranges can be selected in order to verify operation of the EFIS control
panels.
Weather Radar ON/OFF Switch Operational Checks
The EFIS CP's WXR ON/OFF switch should be switched on in order to verify the
display of the magenta, red, yellow, and green sectors which verify the EFIS's
capability for processing and displaying weather radar video data. With the four
weather radar sectors displayed, the control panel's brightness control can be
checked by rotating the inner knob of the EFIS HIS BRT control. Caution should be
exercised when activating the weather radar.
MAP Switch
While the MAP, CTR MAP, or PLAN modes are selected, the various MAP switches on
the EFIS control panel can be actuated to verify that the proper symbol is shown.
During self-test, the RTE DATA switch is inoperative. The VOR/ADF switch causes
the V1, V2, A1, and A2 vectors to be displayed. The NAV AID switch causes the
three VOR, VTC (VOR TACAN), and DME symbols to be displayed. The ARPT switch
causes the single APT (airport) symbol to be displayed. The WPT switch causes a
GRP (ground reference point) symbol to be displayed.
Test Values
All of the data values shown are simulated test values generated within the EFIS SG
and are not dependent upon proper operation of the interfacing systems. Therefore,
the values observed during self-test should coincide with those shown on the
graphic.
TCAS Symbols
If the TFC button is activated on the EFIS control panel, the TCAS symbols shown
on page two will be displayed. TCAS TEST and TRAFFIC will be displayed in all
modes regardless of the position of the TFC button. To properly see these EFISgenerated symbols, it is suggested that a lower range be selected.
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OPERATION
EFIS CONTROL PANEL FAILURES
Failure of the EFIS control panel mode select data to the EFIS symbol generator will cause
the expanded VOR/ILS mode to be displayed on the EHSI. Frequency selection on the VHF NAV
control panel determines whether VOR or ILS information will be displayed.
Loss of range data from the EFIS control panel causes the symbol generator to display range
data from the weather radar transceiver.
Loss of all data, or invalid data from the EFIS control panel will cause both of the above
conditions to be displayed. In addition, the EFIS symbol generator will enable the display
of weather radar data. If the weather radar system is turned on, the data will be
displayed. If the weather radar system is turned off, the message "WXR FAIL" will be
displayed.
With a control panel failure, all TCAS displays to the related EADI and EHSI will be
inhibited. TCAS OFF will be displayed.
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EFIS CONTROL PANEL FAILURE
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SYSTEM TEST
EFIS - SYMBOL GENERATOR FAILURES
Fault Isolation
If any of the conditions occur, which are noted on the graphic, the white "SG FAIL"
message appears on the EADI and EHSI.
During normal operation, the EFIS symbol generator internal display controller rotates
and translates the EADI's airplane symbol from the reference position to the normal
position. In case of a controller failure, this symbol returns to the reference
position, as shown.
If the EFIS symbol generator power supply fails, the EADI and EHSI displays both blank.
Program Pin Parity Faults
At power-up and during self-test, the symbol generator performs a program pin parity
check, including parity program pin DI-54P. If even parity is detected, the SG blanks
both onside display units and displays the message PARITY ERROR and CHECK SG PROGRAM
PIN CONFIGURATION. EFIS BITE remains functional, except self-test displays.
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EFIS SG FAILURES
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NOTES
99
Air Data System
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NOTES
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
PITOT/STATIC SYSTEM
1. Purpose
The purpose of the pitot/static system is to provide a source of dynamic (pitot) and ambient
(static) pressure to airplane systems. The interfacing systems and components convert the
pressure inputs into altitude and airspeed signals.
2. System Description
The pitot/static system is an installation of lines composed of probes, ports, valves,
hoses, tubings, manifolds and various fittings, including drain fittings. The probes sense
dynamic and ambient pressures and supply the pressures through lines to using components.
The probes are heated to prevent ice accumulation.
The alternate static ports provided an alternate source of ambient pressure for the pilot's
pneumatic instruments through static source selector valves in the event of an abnormal
condition existing in the pitot/static probes.
The drain fittings allow the pitot/static system to be drained of accumulated moisture.
PITOT/STATIC SYSTEM
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
PITOT/STATIC SYSTEM COMPONENT LOCATIONS
General Component Locations
Probes and Ports
The pitot/static system consists of: 4 pitot/static probes, 2 alternate static ports,
a window and pitot/static heat module, 12 drain fittings, and assorted equipment; such
as manifolds, tubing, hose and fittings.
Pitot/Static Sources
The pitot/static probes are installed outside the airplane, two on each side, below
control cabin window No. 3 at station 247.6.
The alternate static ports are located at station 406, one at either side of the
airplane.
An alternate static source selector valve is installed on each pilot's sidewall.
Window and Pitot/Static Heat Module
The window and pitot/static heat module is installed on the P5 panel.
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MAINTENANCE TRAINING MANUAL
Pitot Static System - Component Location
EFFECTIVITY
AIRPLANE WITHOUT EFIS
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
PITOT/STATIC DRAIN LOCATIONS
General Component Locations
Five drain fittings are located in the nose wheel well, near the forward bulkhead. The
fittings are:
No.
No.
No.
No.
No.
7
8
9
11
12
(Aux No. 2 pitot)
(Aux No. 2 static)
(F/O static)
(Capt static)
(Aux No. 1 static)
Seven drain fittings are installed on the E-1 rack in the E & E compartment. The fittings
are:
No.
No.
No.
No.
No.
No.
No.
3
4
5
10
13
14
15
(Aux No. 1 pitot)
(Aux No. 1 static)
(Alternate static)
(F/O pitot)
(Capt pitot)
(F/O alternate static)
(Capt alternate static)
Drain fittings No. 1, 2 and 6 have been deleted.
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MAINTENANCE TRAINING MANUAL
Pitot Static System Component Location
EFFICTIVITY
AIRPLANES WITHOUT EFIS
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NOTES
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
PITOT/STATIC PROBE
1. Purpose
The pitot/static probe senses pitot (dynamic) and static (ambient) pressures.
2. Physical Description
The probe consists of three ports: one directed forward for pitot sensing and two on the
side of the probe for sensing static pressure. A strut projects the probe serveral inches
from the airplane skin to minimize effects from airflow turbulence. A baseplate contains the
electrical and pressure fittings. Dowl locator pins, in the baseplate, insure proper
orientation of the probe when installed. The probe is contoured and the ports are located to
minimize errors caused mach number and alpha angle (angle of attack).
An anti-icing heater is installed in the probe to prevent ice accumulation on the probe.
The heater is self-regulating, consuming maximum power during flight and reducing power in
still air. The heater is connected to two shielded pins in the base plate.
CAUTION:
DURING HEATED OPERATION, THE PROBE ATTAINS EXCESSIVE TEMPERATURE. SEVERE BURNS
MAY RESULT WHEN ACCIDENTALLY HANDLED.
PITOT STATIC PROBE
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COMPONENT FUNCTIONAL DESCRIPTION
ALTERNATE STATIC PORT
1. Purpose
The alternate static ports sense static (ambient) pressures and provide the static pressure
input to the alternate static system.
2. Physical Description
The static port is flush-mounted in a cutout in the fuselage skin. A circle is painted
around the port with a caution note below.
4. Caution
The area within the circle must be kept clean and smooth and the holes in the port must not
be deformed or plugged.
3. Access
Access to the static ports is from the forward cargo compartment. Refer to Maintenance
Manual, Chapter 25 for the removal of the cargo compartment lining.
ALTERNATE STATIC PORT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
PITOT/STATIC HEAT MODULE
1. Purpose
The module applies and monitors power to the heaters of the pitot static probes, the TAT
probe, and the alpha vanes.
2. Operation
Placing the "A" switch to the ON position arms the left total air temperature probe heater
and applies power to the upper left and lower left pitot static probes, the left elevator
probe, the left alpha vane.
Placing the "B" switch to the ON position applies power to the lower right and upper right
pitot static probes and the right elevator probe.
When all heater elements draw the proper amount of current, all lights are extinguished. If
the current is too low, the appropriate amber light illuminates. Both master caution lights
and anti-ice warning light will illuminate.
PITOT/STATIC HEAT MODULE
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
STANDBY ALTITUDE/AIRSPEED INDICATOR
1. Purpose
The standby altitude/airspeed indicator provides an alternate source of air data information
for the pilots.
2. Physical Description
The standby airspeed/altimeter indicator is a combination of two flight instruments which
presents an integrated display of indicated airspeed and barometric altitude.
3. Operation
The standby pneumatic airspeed indicator displays indicated airspeed 60 to 450 knots. A
diaphragm within the instrument case is actuated by differential pressure from the airplane
pitot static system. The diaphragm expands or contracts in proportion to the difference
between the pressure in it. This information is transferred from the diaphragm to the drum
counter through a mechanical linkage. The pointer displays indicated airspeed on a drum
graduated in knots.
The standby altimeter is a drum-pointer type having a range of -1000 to 50,000 feet. The
pointer, which indicates feet in hundreds, moves once around the dial for each 1000 feet of
altitude. The drums display feet in thousands and hundreds. Indicators on the face of the
altimeter show barometric pressure in inches of mercury and millibars. The barometric
pressure is set by rotating the knob on the front of the instrument which is adjustable from
948.2 to 1049.5 mb and 28.00 to 30.99 in hg.
An instrument vibrator is provided in the altimeter to shake the instrument to prevent errors
caused by mechanical linkage friction.
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STANBY ALTITUDE/AIRSPEED INDICATOR
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MAINTENANCE TRAINING MANUAL
OPERATION
PITOT/STATIC SYSTEM SCHEMATIC
1. Operation
The schematic shows the interface among components of the pitot/static system and between
the pitot/static system and other airplane systems.
2. Control Sequence
The main sensors of the pitot/static system are four probes, each probe capable of sensing
pitot (dynamic) pressure (P) through the forward facing port of the probe and static
(ambient) pressure through two side ports, S1 and S2. The four probes are identified as the
captain's probe-upper left, the first officer's probe-upper right, the first auxiliary
probe (1st Aux)-lower right, and the second auxiliary probe-lower left. The interface of
the probes through hoses and tubings form four pitot lines and four static lines and are
nomenclatured according to the pitot probe that they are connected to. The S1 and S2 ports
of the upper probes are connected, respectively, to the S1 and S2 ports of the opposite
lower probes to average ambient pressure in the static lines when the airplane is in turns
or during turbulence.
Two aligned static ports are interconnected to provide an averaged alternate sensing of
ambient pressure to the pilot's static source selector valves. A pilot may use the
alternate source for his instruments when ambient pressure sensing from the pitot/static
probes is unreliable.
Twelve drain fittings, connected to the pitot/static lines, allow the lines to be drained of
accumulated moisture.
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Pilot Static System Schematic
EFFECTIVITY
CAL 301-358 PRE-SB 34A1459
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NOTES
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Pilot Static System Schematic
EFFECTIVITY
CAL 301-358 POST-SB 34A1459
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18
RIGHT PROBE HEATER SCHEMATIC
MAINTENANCE TRAINING MANUAL
OPERATION
LEFT PROBE HEATER SCHEMATIC
1. Normal Sequence
The schematic shows the electrical circuits that provide power to the heater elements within the
five probes installed on the left side of the airplane.
Power
Five circuit breakers on the P18-3 panel supply 115 volt ac through segments of the
pitot/static heat switch A and primaries of transformers to heater elements in two
pitot/static probes, one elevator pitot probe, one temperature probe and one alpha vane.
2. Operation
When the switch is closed and normal conditions exist, the heater elements draw current
through the transformers, inducing current in the secondaries and closing S1, S3, S5, S7 and
S9. S2, S4, S6, S8 and S10 do not ground the heat lamps and they remain extinguished.
When the pitot/static heat switch is off, dc from the CAPT PITOT HEAT IND circuit breaker
closes S2, S4, S6, S8 and S10 and the lamps illuminate.
When a heater element opens, the associated lamp illuminates and a ground is applied to the
master caution logic to illuminate the master caution and the anti-ice lights.
The left alpha vane has a vane heater and a case heater. The left alpha vane off light
monitors the vane heater. The case heater is not monitored.
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20
LEFT PROBE HEATER SCHEMATIC
MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
PITOT/STATIC PROBE INSTALLATION
1. Servicing
Probe replacement may be necessary if the probe is physically damaged or if the heater fails.
2. Removal/Installation
The mountng bolts (6) are accessible from the outside of the airplane. Two locating pins are
mounted at opposite corners of the probe mounting flanges
A gasket is used to provide a cabin pressure seal. The gasket is installed between the
probe mounting flange and the airplane structure.
Reference the Maintenance Manual, Chapter 34, Section 11 for complete removal/installation
procedures.
PITOT STATIC PROBE INSTALLATION
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INTRODUCTION
AIR DATA SYSTEM
1. Purpose
The Air Data Computer (ADC) System provides the pilots and interfacing systems with air
pressure related data; such as, altitude, altitude rate and airspeed.
2. System Description
The ADC receives pitot and static air pressure from the pitot probes. It uses the static
pressure to compute the altitude and altitude rate, and the difference between static and
pitot pressure to compute the airplane's airspeed. The computed altitude and airspeed are
then used to compute the mach number.
Electric outputs from the ADC are applied to the flight instruments to provide the pilots
with necessary air data. Signals are provided to the Electric Mach/Airspeed Indicator,
Electric Altimeters, and Electric Vertical Speed Indicators. The electric air data signals
are also applied to other airplane systems.
AIR DATA SYSTEM
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INTRODUCTION
DETERMINING FACTORS FOR COMPUTING AIR DATA INFORMATION
SYSTEM DESCRIPTION
AIR DATA DEFINITIONS
Ps
AMBIENT STATIC PRESSURE:
PRESSURE OF THE STILL AIR SURROUNDING THE
AIRPLANE.
Pt
TOTAL PRESSURE
PRESSURE WHEN THE DYNAMIC PRESSURE IS
BROUGHT TO REST IN A PITOT TUBE.
Qc
DYNAMIC PRESSURE
PRESSURE MEASURED CORRESPONDING TO THE
VELOCITY OF AIRFLOW PAST THE AIRPLANE.
- Ps)
ALT
ALTITUDE
ALTITUDE IS DISTANCE ABOVE SEA LEVEL WHEN
STANDARD PRESSURE CONDITIONS EXIST.
ALTITUDE RATE
ALTITUDE RATE IS VERTICAL SPEED OF THE
AIRPLANE.
MACH NUMBER
MACH NUMBER IS THE RATIO OF TRUE AIRSPEED TO
THE LOCAL SPEED OF SOUND. MACH NUMBER IS ALSO
PROPORTIONAL TO Qc/Ps.
TAT
TOTAL AIR
TEMPERATURE:
STATIC AIR TEMPERATURE PLUS THE
TEMPERATURE RISE CAUSED BY THE AIRFLOW
FRICTIONALLY HEATING AND COMPRESSING THE
TEMPERATURE PROBE.
SAT
STATIC AIR
TEMPERATURE
TEMPERATURE OF THE STILL AIR
SURROUNDING THE AIRPLANE.
IAS
INDICATED AIRSPEED
AIRSPEED PROPORTIONAL TO THE RATIO OF PITOT AND
STATIC PRESSURES.
NO COMPENSATION DONE.
CAS
COMPUTED AIRSPEED
ELECTROMECHANICALLY DERIVED INDICATED AIRSPEED
COMPENSATED FOR STATIC SOURCE ERROR.
TAS
TRUE AIRSPEED
COMPUTED AIRSPEED COMPENSATED FOR CHANGES IN
THE DENSITY AND THE COMPRESSIBILITY OF AIR.
M
23
(Pt
MAINTENANCE TRAINING MANUAL
SIGNAL
DETERMINING FACTOR
STATIC PRESSURE (Ps)
BAROMETRIC PRESSURE
TOTAL PRESSURE
(Pt)
PITOT TUBE PRESSURE
TOTAL PRESSURE
DYNAMIC PRESSURE (Qc)
STATIC PRESSURE
ALTITUDE
(ALT)
STATIC PRESSURE
ALTITUDE RATE
CHANGE IN STATIC PRESSURE
TOTAL PRESSURE
MACH NUMBER (M)
STATIC PRESSURE
TOTAL AIR
Temperature
(TAT)
TEMPERATURE PROBE RESISTANCE
STATIC AIR TEMPERATURE
(SAT)
TEMPERATURE PROBE RESISTANCE MACH
NUMBER
INDICATED AIRSPEED (IAS)
STATIC PRESSURE
TOTAL PRESSURE
COMPUTED AIRSPEED (CAS)
STATIC PRESSURE
TOTAL PRESSURE
TRUE AIRSPEED
(TAS)
TEMPERATURE PROBE RESISTANCE MACH
NUMBER
DETERMINING FACTORS FOR COMPUTING AIR DATA INFORMATION
133763
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NOTES
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INTRODUCTION
DEVELOPMENT OF DIFFERENT AIRSPEEDS
SYSTEM DESCRIPTION
AIRSPEED CORRECTION DEFINITIONS
IN COMPUTING THE DIFFERENT AIRSPEEDS, THE AIR DATA COMPUTER COMPENSATES THE PRESSURE DATA
DEPENDING ON AIRPLANE CONFIGURATION AND FLIGHT CONDITIONS. THE COMPENSATIONS ARE DEFINED BELOW:
STATIC SOURCE ERROR
CORRECTION (SSEC)
CORRECTION FOR STATIC PRESSURE ERRORS
CAUSED BY THE AIRFLOW PAST THE AIRPLANE.
AIR COMPRESSIBILITY
COMPENSATION
CORRECTION FOR THE CHANGE IN
COMPRESSIBILITY OF AIR IN THE PITOT TUBE AS
SPEED AND ALTITUDE VARY.
AIR DENSITY
COMPENSATION
CORRECTION FOR THE CHANGE IN DENSITY
OF AIR AS THE TEMPERATURE AND ALTITUDE VARY.
DEVELOPMENT OF DIFFRENT AIRSPEEDS
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GENERAL DESCRIPTION
AIR DATA SYSTEM COMPONENT LOCATION
General Component Locations
Circuit Breakers
The circuit breakers for ADC-1 are located on the P18 panel, and those for ADC-2 are on
the P6 panel.
Air Data Computers.
The air data computers are located on electronic equipment shelf E1-1.
Autotransformers
The autotransformers are located on the J4 panel in the electronic equipment
compartment.
Indicators
An electric altimeter, vertical speed indicator and mach/airspeed indicator is located
on each of the pilot's panels (P1 and P3). The TAT/SAT/TAS indicator is located on the
center instrument panel (P2).
TAT Probe
The total air temperature probe is mounted on the left side of the airplane, at station
260. The probe is a dual sensor with sensor 1 feeding ADC-1 and sensor 2 feeding ADC2.
Mach/Airspeed Warning Clackers
The Mach/Airspeed warning clackers are located inside the aural warning devices unit.
The aural warning devices unit is mounted on the forward right side of the control
stand.
Mach-Airspeed Warning Test Switches
The Mach/Airspeed Warning Test Switches one for each system, are located on the aft
overhead panel (P5).
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AIR DATA SYSTEM COMPONENT LOCATION
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COMPONENT FUNCTIONAL DESCRIPTION
AIR DATA COMPUTER
1. Purpose
The ADC system provides the electrical air data information required by the pilots' flight
instruments and by other airplane systems.
2. Physical Description
The ADC is a rack mounted assembly and uses two camlock hold down devices for retension in the
shelf.
The ADC assembly consists of: a chassis assembly, power supply, 12 plug-in circuit cards,
and two pressure transducers.
The front panel provides two switches for self-test actuation, a self-test indicator light, a
test connector for use during shop testing of the ADC, and two pneumatic connectors for
connection to the airplane pitot and static system. Three electrical connectors, on the
rear of the ADC, provide the electrical interface with the airplane wiring.
3. Operation
The ADC, using digital and microprocessor techniques, processes pitot and static pressure
and total air temperature to produce the required air data information. The ADC provides:
altitude, altitude rate, vertical speed, airspeed, mach, true airspeed, total air
temperature, and static air temperature to the various indicators and systems.
The air data signals are provided in analog form and by ARINC 429 buses. Parallel digital
altitude is provided for the ATC systems.
4. Self-Test
The TEST SELECT switch manually selects built-in test to be performed as follows:
FUNCTION - transmits fixed values on all the computer signal output lines.
SLEW - sets altitude rate to 600 feet per minute.
FAIL - activates all failure warning flags.
The PUSH TO TEST pushbutton switch initiates the built-in test. The TEST VALID WHEN LIT
indicator light comes on within 2 seconds after the test is successfully completed and
remains on when there is no failure.
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AIR DATA COMPUTER
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COMPONENT FUNCTIONAL DESCRIPTION
ELECTRIC ALTIMETER
1. Purpose
The altimeter displays the barometric altitude of the airplane. It also contains controls
for setting the barometric correction, and an altitude reference.
2. Physical Description
The altimeter is a round, panel-mounted instrument. The face of the instrument contains a
digital and an analog display of altitude, a digital display of the manually set barometric
pressure reference, an altitude reference cursor, and controls to position the BARO set
indicator and the altitude reference cursor.
3. Power
The altimeter requires 26 volts ac power. It contains a power supply to provide the
internally required dc voltage.
4. Operation
The altimeter receives altitude data from the ADC. An electro-mechanical servo output,
mechanically summed with barometric correction, positions the digital altitude indicator and
the analog altitude pointer.
The digital altitude display indicates between -1000 ft. and 50,000 ft. When the altitude is
below sea level, a NEG flag covers the first two digits. An OFF flag covers the first digit
when the ADC or the altimeter malfunctions, or when power is removed from the ADC system.
The altitude pointer makes one complete revolution for each 1000 ft. change in altitude.
The pointer's 1000 ft. circular scale has 20 ft. increments.
The BARO set indicator is displays pressure in inches of mercury (IN HG) and millibars (MB).
The BARO set knob is used to set the indicator to the local sea level pressure or to the
standard sea level pressure of 29.92.
When the indicator is set to 29.92, the altitude indication is pressure altitude, and when
set to the ambient pressure, the altitude indication is BARO-corrected altitude.
The altitude reference knob is used to mechanically position the altitude reference cursor.
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ELECTRIC ALTIMETER
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COMPONENT FUNCITONAL DESCRIPTION
VERTICAL SPEED INDICATOR
1. Purpose
The vertical speed indicator provides a visual indication of the rate of climb or descent of
the airplane.
2. Physical Description
The vertical speed indicator is a round, panel-mounted instrument. Its features are: a
pointer indicating altitude rate when read against a vertical speed scale, vertical speed
scale, and an OFF flag with fire orange lettering on a white background. When the flag is
retracted, the aperture appears black. Edge lights illuminate the face.
3. Power
The vertical speed indicator requires 115volts ac power. It contains a power supply that
provides the required dc voltage.
4. Operation
The indicator receives ADC altitude rate. A servo positions the vertical speed pointer.
Full scale indication is 6000 ft/min, with the upper half of the scale indicating climb and
the lower half descent.
The OFF flag is out of view when the system is functioning properly; and is in view when
power is off or with an ADC or indicator malfunction.
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RA/VSI INDICATOR
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COMPONENT FUNCTIONAL DESCRIPTION
ELECTRIC MACH AIRSPEED INDICATOR
1. Purpose
The Mach Airspeed Indicator provides visual displays of airspeed, airspeed limit, mach
number, and target airspeed. It provides the means to manually select the target airspeed.
The indicator also provides the control circuitry for the mach/airspeed aural warning.
2. Physical Description
The Mach Airspeed Indicator is a round, panel-mounted, instrument. Its' features are: a
digital airspeed display, an analog airspeed pointer, an orange and white striped maximum
airspeed pointer, a target airspeed cursor, and a digital mach number indicator. A speed
reference knob is provided to manually position the target airspeed cursor.
The airspeed range is from 60 to 450 knots CAS, and the mach range is from 0.400 to 0.999
mach. The analog airspeed scale is logrithmic from 60 to 250 knots, and linear from 250 to
450 knots.
Failure flags are provided to indicate malfunction of the airspeed display, mach display,
target airspeed cursor, and the maximum airspeed (Vmo) pointer. A flag is provided to
annunciate manual operation of the target airspeed cursor.
Several external index markers are provided that are manually positioned around the
periphery of the airspeed scale.
3. Power
The Mach Airspeed Indicator requires 26 volt ac power. An internal power supply provides
the necessary dc voltage.
4. Operation
The Mach Airspeed Indicator receives airspeed and altitude data from the ADC, and target
airspeed from the Flight Management Computer System. An integral computer calculates the
mach number and maximum airspeed, and controls the maximum airspeed warning.
The digital airspeed display and the analog airspeed pointer are positioned by a servo that
receives the airspeed signal from the ADC. The digital mach display and the maximum
airspeed pointer are positioned by separate computer controlled servos. When the speed
reference knob is pushed in, the target airspeed cursor is positioned by a servo. When the
knob is pulled out, the servo is disabled and the target airspeed cursor is manually
positioned through the speed reference knob. Manual control is annunciated by the manual
mode flag.
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COMPONENT FUNCTIONAL DESCRIPTION
4. Operation (Cont)
The computer closes a switch that is used to enable the mach/airspeed warning when the
airspeed equals or exceeds the computed maximum airspeed. Maximum airspeed is depicted by
the graph. Below 25,968 ft., the airplane is airspeed limited (Vmo) to 340 knots CAS.
Above that altitude, the airplane is mach limited (Mmo) to 0.82 mach; and when mach is held
constant, the maximum airspeed decreases as altitude increases. The airplane is altitude
limited to 37,000 ft.
Airspeed, mach, Vmo, or airspeed cursor flags will be in view when power is removed from the
ADC system or if there is an associated failure in the ADC or indicator. The airspeed
cursor INOP flag is mechanically retracted when the speed reference knob is pulled out.
The Mach Airspeed Indicator has provisions for sending a speed error signal to the
autothrottle. The signal is the difference between the displayed airspeed and the target
airspeed.
ELECTRIC MACH AIRSPEED INDICATOR
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NOTES
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Mach Airspeed Warning System Schematic
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
AURAL WARNING DEVICES UNIT
1. Purpose
The Aural Warning Device Unit provides sounds warning of conditions that require immediate
attention.
2. Physical Description
The Aural Warning Devices Unit is 10 in. long, 5 in. wide, and 2 in. deep. It is mounted
vertically on the aft end of the forward electronics panel (P9). Two connectors, located on
the rear of the unit, provide the interface with the airplane wiring.
The unit contains two clackers that provide the mach airspeed warning, the fire warning
bell, and a speaker that emits warning sounds pertaining to takeoff warning, pressurization
warning, landing warning, and selcal.
3. Operation
The clacker devices produce the Mach/Airspeed Warning sound. The sound is distinctly
different than any other sound in the airplane, and is easily identified. Each clacker is
provided 28 volt dc and is enabled by a ground from the Mach Airspeed Indicators, when an
overspeed condition is detected. Clacker-1 interfaces with the captain's Mach Airspeed
Indicator, and clacker-2 interfaces with the first officer's indicator.
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AURAL WARNING DEVICES UNIT
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
TOTAL AIR TEMPERATURE PROBE
1. Purpose
The total air temperature (TAT) probe senses the air temperature affecting the performance
of the airplane and converts the sensing into analog signals.
2. Physical Description/Features
The TAT probe is a small metal strut, secured to the airplane by six screws. It allows air,
entering a port on the front of the probe, to flow across temperature sensing elements, and
to exit through ports on the side and aft surface of the strut. The TAT probe is installed
with the inlet probe facing forward.
The temperature sensitive wire elements are hermetically sealed in concentric tubes.
Airflow around the elements causes the resistance to vary as a function of the total air
temperature. Air must flow through the probe for proper operation. The TAT probe contains a
heating element for de-icing. The probe design allows heated air to flow out of the probe
without affecting the sensing elements. The TAT probe is installed with the inlet probe
facing forward.
TOTAL AIR TEMPERATURE PROBE
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
TAT/SAT/TAS INDICATOR
1. Purpose
The TAT/SAT/TAS indicator provides the display of total air temperature (TAT) from -99°C to
+50°C, static air temperature (SAT) from -100°C to +50°C, and true airspeed (TAS) from 100
knots to 599 knots. The data is received from the ADC on an ARINC 429 digital data bus.
2. Physical Description/Features
The indicator is a round, panel-mounted instrument. It features a three-digit, white
on black background, liquid crystal display (LCD), and a push-push-push function selector
switch.
The indicator requires 26 volt ac input power, and contains a power supply to provide the
required dc voltage. Electrical power for edge lighting of the LCD is provided from the
airplane panel lighting bright/dim circuit.
TAT is the initial mode when power is applied. To step to the next mode, press the function
selector push button. The mode sequence is TAT, SAT, TAS, and then return to TAT.
The display is tested through the master dim and test circuit. The test input causes the
display to alternate between, all segments ON for two seconds, and blank for one second.
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TAT/SAT/TAS/INDICATOR
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OPERATION
AIR DATA SYSTEM 1-BLOCK DIAGRAM
Operation Sequence
Air data system 1 receives pitot and static pressure from the captain's pitot and static
sources, total air temperature from a TAT probe, and baro-correction from the
captain's and first officer's altimeters. The air data computer (ADC), using these inputs,
computes: altitude, altitude rate, computed airspeed (CAS), mach (M), true airspeed (TAS),
total air temperature (TAT), and static air temperature (SAT). These air data measurements
are distributed as synchro signals, potentiometer signals, parallel digital data, and as
serial digital data on ARINC 429 digital data buses.
Altitude Outputs
Coarse and fine synchro altitude is distributed to the captain's altimeter and
mach airspeed indicators and the flight data recorder. An altitude signal is
supplied to the cabin pressure controller. Altitude rate is provided to the
captain's vertical speed indicator and the ground proximity warning system.
Digital altitude data is provided to the ATC transponder. The captain's
altimeter provides baro-correction to the cabin pressure controller.
Synchro CAS is distributed to the captain's mach airspeed indicator, the autothrottle
computer, auto slat computer 1, and the flight data recorder. An airspeed switch
provides a bi-level gain control for the yaw damper. Synchro mach is provided to the
ground proximity warning system.
ARINC 429 Bus Outputs
There are four digital data buses providing distribution of the air data measurements.
Digital bus 1 supplies FCC-A with pressure altitude, baro-correction, mach, CAS, TAS,
SAT, TAT, and altitude rate.
Digital bus 2 provides the TAT/SAT/TAS indicator with TAT, SAT, and TAS.
Digital bus 3 provides the autothrottle computer with: pressure altitude, barocorrection, mach, CAS, TAT, SAT, and altitude rate; and provides the stall warning
computer-1 with computed airspeed.
Digital bus 4 provides the flight management computer with pressure altitude, barocorrection, mach, CAS, TAS, TAT, SAT, and altitude rate and, the IRS-1 with pressure
altitude, altitude rate, and TAS.
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MAINTENANCE TRAINING MANUAL
Air Data System No. 1 Block Diagram
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MAINTENANCE TRAINING MANUAL
OPERATION
AIR DATA SYSTEM 2-BLOCK DIAGRAM
Operation Sequence
Air data system 2 receives pitot and static pressure from the first officer's pitot and
static sources, total air temperature from a TAT probe, a n d b a r o - correction from the
captain's and first officer's altimeters. The air data computer (ADC), using these inputs,
computes altitude, altitude rate, computed airspeed (CAS), mach (M), true airspeed (TAS),
total air temperature (TAT), and static air temperature (SAT). The air data measurements
are distributed as synchro signals, potentiometer signals, parallel digital data, and as
serial digital data on ARINC 429 digital data buses.
Altitude Outputs
Coarse and fine synchro altitude is provided to the first officer's altimeter
and mach airspeed indicators. Altitude rate is provided to the first officer's
vertical speed indicator. Digital altitude data is provided to the ATC
transponder. The first officer's altimeter provides baro-correction to the
cabin pressure controller.
Airspeed Outputs
Synchro CAS is distributed to the first officer's mach airspeed indicator and auto slat
computer 2.
ARINC 429 Bus Outputs
There are four digital data buses to provide distribution of the air data measurements.
Digital bus 1 supplies FCC-B with pressure altitude, baro-correction, mach, CAS, TAS, SAT,
TAT, and altitude rate.
Digital bus 2 provides the TAT/SAT/TAS indicator with TAT, SAT, and TAS.
Digital bus 3 provides the autothrottle computer with pressure altitude, baro-correction,
mach,CAS, TAT, SAT, and altitude rate and the stall warning computer-2 with computed
airspeed.
Digital bus 4 provides the flight management computer with pressure altitude, barocorrection, mach, CAS, TAS, TAT, SAT, and altitude rate and IRS-2 with pressure altitude,
altitude rate, and TAS.
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
ADC SELF-TEST
Functional Test
There are three different routines in the ADC self-test. They are FUNCTION, SLEW, and FAIL.
Each test routine is selected through the test selector switch. The selected routine is
initiated by pressing and holding the PUSH-TO-TEST switch. The switch is held until the
appropriate test responses have been observed. Illumination of the (Red) TEST LIGHT signifies
a valid evaluation of the test, by the ADC's self-test circuit. An assistant is required to
verify correct response of the instruments and indicators, to the ADC signal outputs.
Test Evaluation
With a valid ADC, the TEST LIGHT will illuminate after about two seconds during the FUNCTION
test; and immediately, during the SLEW and FAIL tests.
Allow the instruments to stabilize, before evaluating their response to the ADC signal
outputs (about 45 seconds during FUNCTION).
During the ramping of the ADC output signals, and stabilization of the instruments, the
warning flags are erratic. Do not evaluate their response until after the instrument has
stabilized.
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TEST SELECT
SWITCH.POSITION
FUNCTION
PUSH TO TEST
SWITCH POSITION
PRESS AND HOLD FOR
45 SEC
TEST LIGHT
STATUS
ILLUMINATED
SIGNAL OUTPUTS
WAIT APPROXIMATELY 30 SECONDS SECBEFORE
VERIFYING THE FOLLOWING VALUES ON THE
INDICATORS
ALT = 10,000 ±40 FT
MACH = 0.785 ±0.010 M
CAS = 440 ±4NM
TAS = 477.8 ±6 KTS
SAT = -29.4 ±2ºC
TAT = 0 ±2ºC
NO FLAGS
SLEW
PRESS AND HOLD FOR
10 SEC
ILLUMINATED
ALT RATE = -600 ±30 FPM
MASTER FAILURE WARNING VALID
AND FAILURE DISCRETES OUTPUT
ARE VALID
FAIL
PRESS AND HOLD FOR
15 SEC
ILLUMINATED
MASTER FAILURE WARNING & FAILURE
DISCRETE OUTPUTS ARE FAILED.
HONEYWELL HG480B40 SERIES
ADC SELF TEST
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NOTES
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EFIS Differences
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Pilot Static System Component Location
EFFECTIVITY
AIRPLANES WITH EFIS
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Pilot Static System Schematic
Figure 2
EFFECTIVITY
AIRPLANES WITH EFIS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
AIR DATA COMPUTER
Purpose
The Air Data Computer (ADC) converts inputs from the pitot static system and the Total Air
Temperature (TAT) Probe into air data information for primary flight displays, navigation
and altitude reporting.
Physical Description
The ADC is enclosed in a 1/2 Austin Trumbull Radio (ATR) long case.
On the front of the unit, there are two pneumatic connectors, a toggle switch (for checking
current status or past history data) and a Light Emitting Diode (LED) display window (for
showing test results).
Operation
The ADC uses a digital microprocessor to process pitot pressure, static pressure and total
air temperature to produce the required air data information. The ADC provides altitude,
altitude rate, baro corrected altitude, computed airspeed, mach, true airspeed, total air
temperature, static air temperature, and total and static air pressures, to the various
indicators and systems.
The air data signals are sent to other systems in analog and digital formats. The digital
information is transmitted using Aeronautical Radio Incorporated (ARINC) 429 buses.
Self-Test
The TEST/HISTORY switch allows selection of built-in-test. TEST selection checks current
status of the ADC and associated inputs. HISTORY selection checks failure information for
up to ten previous flights.
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AIR DATA COMPUTER
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COMPONENT FUNCTIONAL DESCRIPTION
EADI-AIR DATA DISPLAYS
Purpose
The Electronic Attitude Director Indicator (EADI) displays computed airspeed and mach from
the air data system.
Operation
The Electronic Flight Instrument System (EFIS) Symbol Generator receives mach, computed
airspeed, and true airspeed from the Air Data Computer (ADC). The symbol generator then
provides the video signals to drive the air data displays on the EADI.
In-Flight Normal Displays
The airspeed display consists of a calibrated white airspeed tape and a numerical readout of
current airspeed. The airspeed tape displays a range of 84 knots with indices every 10
knots and numeric displays every 20 knots.
Mach is displayed numerically in white numbers below the tape. The display appears when mach
is greater than or equal to 0.40, and is removed when mach is less than 0.38.
On-Ground Normal Displays
On the ground, with the airplane stationary, the Air Data Computer provides a minimum
computed airspeed of 45 knots. This value is displayed on the speed tape. The mach display
is blank below 0.40 mach.
Invalid Displays
When the airspeed or mach data word from the Air Data Computer is invalid, or when in the
air and the computed airspeed is less than 50 knots, the corresponding display blanks and a
yellow flag appears.
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DATA DISPLAYS - ELECTRONIC INSTRUMENTS
MAINTENANCE TRAINING MANUAL
OPERATION
AIR DATA SYSTEM BLOCK DIAGRAM
Operational Sequence
The air data systems receive pitot and static pressure from their respective pitot-static
sources, total air temperature from the total air temperature probe, and barometric
correction from the altimeters. The air data computer provides altitude, altitude rate,
computed airspeed, Mach, true airspeed, total air temperature, and static air temperature.
These outputs are distributed as synchro signals, potentiometer signals, and as serial
digital data on four, identical Aeronautical Radio Incorporated (ARINC) 429 digital data
buses.
Analog and Discrete Signals
Coarse and fine synchro altitude signals are distributed to the onside altimeter and
Mach airspeed indicator from the onside air data computer. An altitude potentiometer
signal from air data computer-1 and the barometric correction signal from the captain's
and first officer's altimeters are sent to the cabin pressure controller.
Synchro computed airspeed is sent to the onside Mach airspeed indicator from the onside
air data computer. Air data computer-1 sends a computed airspeed synchro signal to the
autothrottle system and a discrete gain control signal to the yaw damper system.
Overspeed discrete signals are sent from the Mach airspeed indicator to the aural
warning module if the maximum operating airspeed (Vmo) or maximum operating mach (Mao)
is exceeded. A test discrete from the mach airspeed test module is received by either
indicator to test the aural warning.
ARINC 429 Data Bus Outputs
Each air data computer has four, identical digital data buses which provide
distribution of airdata information.
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AIR DATA SYSTEM BLOCK DIAGRAM
MAINTENANCE TRAINING MANUAL
SYSTEM TEST
AIR DATA COMPUTER SELF-TEST
Two tests may be performed using the TEST/HISTORY switch on the front panel of the air data
computer (ADC).
TEST
Momentarily placing the TEST/HISTORY switch in the TEST position performs a self-test of the
air data computer. The hexadecimal display is tested and the display "8 8" will appear if
the display is operating properly. After the display test, a display of "A A" indicates
that the self-test is occurring and the display values on the graphic will be displayed in
the flight compartment. The output test values will remain for approximately sixty seconds
and the display will flash for the last ten seconds of the test. The operator may select
TEST again during this ten second period to have the value displayed for another sixty
seconds.
If failures are detected in the ADC or with sensor input signals, the display will
momentarily show "F F" followed by the faults displayed as "F X" where the "X" will be a
digit from 1 to 9. The meaning of the fault codes is defined on the front of the ADC. If
more than one fault exists, the faults will be displayed in order. BARO REF #1 for Air Data
Computer 1 is the captain's BARO reference while BARO REF #2 is the first officer's. BARO
REF #1 for Air Data Computer 2 is the first officer's BARO reference while BARO REF #2 is
the captain's.
History
The ADC stores in nonvolatile memory any faults that occur during a flight segment for the
last ten flights.
Momentarily placing the TEST/HISTORY switch to HISTORY will cause the display to show any
faults that occurred during the last flight. This information will be displayed as "0 X"
where "0" is the flight number (last flight) and "X" is the fault code. All faults for that
flight will be displayed in order. Preceding flights may be selected by again placing the
switch to HISTORY.
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AIR DATA COMPUTER (ADC) SELF-TEST
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INTRODUCTION
AURAL WARNING SYSTEM
Purpose
The flight compartment aural warning system provides characteristic audio signals to alert
the flight crew to: an abnormal takeoff condition, landing condition, autopilot disengage,
pressurization condition, Mach airspeed condition, engine fire, APU fire, wheel well fire,
selcal alert, or crew calls.
System Description
The Aural Warning System consists of separate system monitor circuits and an aural warning
devices unit. The monitor circuits interface with the airplane systems to detect certain
potentially dangerous flight control configurations, critical system warnings, and flight
crew alerts. The aural warning devices unit interprets signals from the monitor circuits
and provides audio alert signals to the flight crew.
The systems which interface with the Aural Warning System include:
Landing Gear Warning System
Takeoff Warning System
Cabin Altitude Warning System
Digital Flight Control System
Mach Airspeed Warning System
Fire Detection System
Crew Call System
Selcal System/ACARS
The specific conditions which initiate an aural warning alert for each of these systems are
covered in the individual sections and will not be reviewed here.
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AURAL WARNING SYSTEM
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COMPONENT LOCATION
AURAL WARNING SYSTEM COMPONENT LOCATION
General Component Locations
Aural Warning Devices Unit
The Aural Warning Devices Unit is located on the aft right side of the forward
electronics panel (P9).
Circuit Breakers
Four circuit breakers provide necessary voltages to the Aural Warning Devices Unit.
Three circuit breakers are located on the P-6 panel and one circuit breaker is located
on the P-18 panel.
AURAL WARNING SYSTEM COMPONENT LOCATIONS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
AURAL WARNING UNIT
Purpose
The aural warning unit produces several different sounds; any one of which alerts the crew
to conditions that require immediate attention.
Physical Description
The unit is 10 inches long, 5 inches wide, 2 inches deep and is mounted vertically on the
aft right side of the P9 panel. Two electrical connectors at the rear of the unit provide
interface with the airplane wiring.
BITE Switch
There is a BITE switch located on the top of the aural warning unit. Turning the switch to
the left tests channel A and turning the switch to the right tests channel B. The wailer
and horn sounds are heard when the switch is turned while the clacker sound is heard when
the switch is released.
AURAL WARMING UNIT
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NOTES
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OPERATION
AURAL WARNING SYSTEM SCHEMATIC
Power
The aural warning unit receives 28 volt dc power from four circuit breakers. The 28 volt
BAT BUS is connected through the Master Warning and Control circuit breaker and the Aural
Warning circuit breaker on the P6 panel. The master warning and control circuit powers the
bell sound and the aural warning circuit powers the horn, chime and wailer sounds. The 28
volt ELEX PWR-1 is connected through the Aural Mach and Airspeed Warning circuit breakers.
One circuit breaker is on the P6 Panel and one is on the P18 panel. The Aural Mach and
Airspeed Warning circuits power the clacker sounds.
There are two identical channels in the aural warning unit. Both channel circuits are
powered by the four sources of 28 volt power.
Operation
The aural warning unit channels are wired in parallel and are redundant. If one channel
fails, the flight crew will notice a 6 db loss in volume from the aural warning unit.
Each channel contains a power supply, a ground-activated input circuit, a power (28 volts
dc) activated input circuit, a controller, an aural synthesizer, a speaker driver and a
speaker.
Ground-Activated Input Circuit
The ground inputs and the sounds associated with them are: the fire warning input - bell
sound, two overspeed warning inputs - clacker sound, the takeoff warning input and cabin
pressure warning input - intermittent horn sound, the landing warning input - steady
horn sound, and the autoflight disconnect input - wailer sound.
Power-Activated Input Circuit
The power discrete inputs consists of a 28 volt signal when the sound is to be
activated. The power inputs and the sounds associated with them are: the crew call
input - high chime sound and the (optional) selective calling (SELCAL) or ACARS input high/low chime sounds.
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OPERATION
Operation (Cont.)
Controller
The controller receives signals from both input circuits and prioritizes the sound
generation of the horn sounds and the wailer sound. An intermittent horn sound has the
highest priority, the steady horn sound has second priority, and the wailer sound has
third priority. The other sounds are not prioritized and may be present at the same
time as the horn or wailer sounds.
After prioritizing the sound generation, the controller causes the aural synthesizer to
emit the correct sound or sounds.
Aural Synthesizer
The aural synthesizer is capable of emitting all the sounds required of the aural
warning unit. It is also capable of emitting multiple sounds at the same time.
Speaker Amplifier and Speaker
The signals from the aural synthesizer are amplified by the speaker amplifier and sent
to the speaker.
BITE Switch
The BITE switch is located on the top of the aural warning unit. Use a screwdriver to
turn the switch. Channel A is tested by turning the switch to the left and channel B
by turning to the right. The horn and wailer sounds will be heard when the switch is
turned and the clacker sound will be heard when the switch is released.
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AURAL WARNING SYSTEM SHEMATIC
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MAINTENANCE PRACTICES
AIR DATA SYSTEM - MAINTENANCE SUMMARY
Operational Checkout
Refer to the Maintenance Manual Chapter 34 Section 12 for complete troubleshooting, testing
and component removal or installation procedures.
Precautions
The pitot static probes, alternate static ports and temperature probes may be heated to
prevent ice formation. Consequently, these devices can become extremely hot. Before
handling them, assure that the heating power is o f f a n d that the devices are cool.
Damage to system components can occur during testing procedures if excessive pitot pressure
or static vacuum is applied. If the static pressure exceeds the pitot pressure by more than
10 inches of mercury, the ADC will likely be damaged.
Any test which causes the mach airspeed clacker to sound should be limited to 5 minutes to
avoid overheating the clacker. Allow a minimum cooling period of 15 minutes before engaging
the clacker again.
Power
Prior to operating, testing or troubleshooting the ADC system, assure that the proper power
is available and all related circuit breakers are set.
Tests
A complete self-test of the system is available at the front panel of the ADC. Consult the
Maintenance Manual for particulars.
Removal/Replacement
The ADC may be damaged if it is not at ambient pressure when removed or installed.
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MAINTENANCE TRAINING MANUAL
PRECAUTIONS
ASSURE PITOT STATIC PROBE, ALTERNATE
STATIC PORTS AND TEMPERATURE PROBES
ARE COOL BEFORE HANDLING.
AVOID APPLICATION OF EXCESSIVE PRESSURE
OR VACUUM TO THE SYSTEM.
AVOID PROLONGED OPERATION OF THE MACH/ AIRSPEED
WARNING CLACKER.
POWER
BEFORE OPERATING,
ASSURE POWER IS APPLIED TO ADC,
INSTRUMENTS AND RELATED SYSTEMS.
TEST
EXTENSIVE SELF TEST IS AVAILABLE
MACH/AIRSPEED WARNING TEST IS PERFORMED
IN THE FLIGHT COMPARTMENT
REMOVAL/INSTALLATION
ASSURE ADC IS AT AMBIENT PRESSURE PRIOR
TO REMOVAL.
108326
AIR DATA SYSTEM – MAINTENANCE SUMMARY
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NOTES
72
Inertial Reference
System
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
INERTIAL REFERENCE SYSTEM
1. Purpose
The inertial reference system (IRS) is one part of the FMS.
The Honeywell inertial reference system is a laser gyro and accelerometer based reference
system. Among the many parameter outputs, the IRS provides attitude, heading, acceleration
and angular outputs, together with ground speed, drift angle and present position
information.
2. System Description
To perform its function, the IRS consists of two inertial reference units (IRU), a common
mode selector unit (MSU) and a common inertial system display unit (ISDU) and two digital to
analog adapters (DAA).
The MSU controls system mode selection and the ISDU provides operator interface to the
system. The IRUs provide bus information to the ISDU, the flight management computer (FMC),
the flight control computer (FCC), the autothrottle computer (A/T), the DAAs and to the
pilots' horizontal situation indicators (HSI).
The DAAs provide attitude and heading information to the pilots' instruments. The IRUs are
provided with bus information from the ISDU, the air data computers (ADC) and the FMC.
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INERTIAL REFERENCE SYSTEM
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
IRS COMPONENT LOCATIONS
General Component Locations
Circuit Breakers
The circuit breakers for supplying ac and dc power to the IRUs and the DAAs are located
on the P18-2 and P6-1 circuit breaker panels. Ac circuit breakers for the pilot's
instruments are also located on P18-2 and P6-1. Dc power for the transfer relays comes
from a single circuit breaker on P18-2.
Inertial Reference Units
The IRUs are mounted on a factory aligned shelf, E3-5, in the electrical/electronic
compartment.
Digital Analog Adapters
The DAAs, together with the flight instrument accessory unit, the attitude transfer
relays 1 and 2 and the compass transfer relays 1 and 2, are mounted on the E1-2 shelf in
the electrical/electronic compartment.
Control Units
The ISDU and the MSU are mounted on the aft overhead panel P5. The flight management
computer system control display units (FMCS-CDU) are mounted on the forward electronic
panel P9.
Instruments
An ADI, an HSI and a RDDMI are mounted on each pilot's panel, P1 and P3. The
instrument transfer panel, which contains the compass and the attitude transfer
switches, is mounted on the forward overhead panel, P5.
IRS Master Caution Unit
The master caution unit is mounted on the P10 panel on the right side of the flight
control cabin.
IRS Master Caution Light
The IRS master caution light is mounted on the left side of the pilots' glareshield,
P7.
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MAINTENANCE TRAINING MANUAL
IRS COMPONENT LOCATIONS
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NOTES
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
IRS MODE SELECT UNIT
1. Purpose
The mode select unit (MSU) provides mode selection for two IRUs through two mode select
switches and monitors the operation of the IRUs through two sets of alert lights mounted on
the unit.
2. Location
The MSU is mounted on the aft overhead panel P5 in the flight control cabin.
3. Physical Description
The MSU contains two, three stack, four position mode select switches, one for the left (L)
IRU and one for the right (R) IRU. The switches are detented in each position. The
positions are OFF, ALIGN, NAV, and ATT.
Two sets of four light switch assemblies are installed on the unit, one light for an alert
purpose and three for warning purposes. The white alert light is the ALIGN annunciator and
the three amber warning lights are the ON DC, FAULT and DC FAIL annunciators.
The MSU is mounted with four DZUS fasteners and is interfaced with airplane wiring through
two rear mounted connectors.
4. Power
Dc power for lamp illumination is derived from the IRU or from the master dim test circuit.
5. Operation
When the mode selector switches are in OFF, power is removed from the IRUs.
A.
Align
When ALIGN is selected, power is applied to the IRUs and the IRUs normally progress
through an alignment period of approximately ten minutes before the navigational mode
is armed. When the switches are maintained in ALIGN, however, the IRUs remain in the
align mode. Normal alignment requires the entry of present position into the IRUs.
B.
Navigation
The NAV position of the switches enables the navigational mode, provided the alignment
period is completed in the IRUs.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
5. Operation
B.
Navigation (Cont.)
The NAV position may be selected directly from OFF. The NAV position is detented and
requires pulling the knob outward for placement out of NAV.
C.
Attitude
The ATT position of the switches places the IRUs into a reversionary mode of operation.
6. Monitor
A.
Align
The ALIGN light illuminates steady during alignment or flashes when an abnormal
alignment is sensed.
B.
On DC
The ON DC lights illuminate when 115 volt ac power is removed from the IRUs and the
IRUs operate on 28 volt dc. The 28 volt dc continues to power the left IRU but after
five minutes, power is removed from the right IRU.
C.
Fault
The FAULT lights illuminate when an abnormal condition exists in the IRUs.
D.
DC Fail
The DC FAIL lights illuminate when the airplane battery power is insufficient to
maintain IRU operation.
7. Lamp Test
The lights may be tested by pressing each light switch assembly, by activating the brightdim-test circuit or by initiating an IRU test.
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IRS MODE SELECT UNIT
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MAINTENANCE TRAINING MANUAL
OPERATION
MODE SEQUENCES
1. Normal Sequence
A.
Turn On
The normal sequence of modes in the IRS is to switch from OFF to either ALIGN or NAV.
This action starts the alignment sequence. During alignment the IRU establishes local
vertical, true heading, and latitude. This requires approximately 10 minutes. Present
position is entered during this time. If the switch is placed to NAV, the IRU will
automatically sequence to NAV at the end of the alignment sequence.
B.
Downmode Align
The downmode to align is activated by rotating the IRU mode select switch from NAV to
ALIGN position providing the ground speed is not greater than 20 knots. This action
should normally be performed on intermediate short stops when the IRS is left operating
and insufficient time is available to perform a complete 10-minute alignment to correct
existing navigation errors. The procedure can also be used to trim errors during
extended waiting periods for takeoff or during delays prior to departure.
The duration of the align downmode is 30 seconds following activation or whenever the
mode select switch is rotated to the NAV position, whichever is greater.
The result of this action is to remove the velocity errors (set groundspeed to zero)
and to re-erect the navigation reference coordinate frame to level (remove pitch/roll
attitude errors). Errors is position can also be corrected if the option to re-enter
ramp latitude and longitude is also exercised.
If the IRU is allowed to remain in the align mode for a period greater than 30 seconds,
the alignment filter will begin to correct airplane heading. However, the difference
between this procedure and the normal align is that NAV mode can be entered at any
time and the normal 10-minute waiting period is not required. In this situation the
degree of refinement of the prior NAV mode heading will be a function of the operating
time in the align mode.
C.
Power Down
The normal power down sequence occurs any time the switch is moved to OFF. The ALIGN
annunciator illuminates for 30 seconds. During this 30-second period, all in-flight
faults detected by the IRU are stored in non-volatile memory.
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IRS Modes
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
INERTIAL SYSTEM DISPLAY UNIT
1. Purpose
The inertial system display unit (ISDU) provides pilot interface with the IRUs. The ISDU
allows entry of initialization data for the IRUs. Displays of track angle, ground speed,
present position, wind direction and speed, heading and system status are available.
2. Location
The ISDU is located on the aft overhead panel, P5.
3. Physical Description
The ISDU front panel contains two seven segmented displays. Mounted on the left side of the
front panel are a five position display select (DSPL SEL) switch, including a brightness
control, and a two position system display (SYS DSPL) switch. On the right side of the
panel are a set of twelve keys called the keyboard. The unit weighs four pounds (1.8 kgms)
and is interfaced with airplane wiring through two rear mounted connectors.
4. Power
The ISDU is powered by a 28 volt dc source from each of the IRUs.
5. Control
A.
Display Selector Switch (DSPL SEL)
The DSPL SEL switch controls the navigational data displayed on the left and right
displays:
TK/GS (Track Angle/Ground Speed) - True track angle from 0 to 359.9 degrees is
displayed in the left display with a resolution of 0.1 degree. Ground speed from 0 to
2000 knots is displayed in the right display with a resolution of 1 knot.
PPOS (Present Position) - Latitude from 90°S to 90°N is displayed in the left display
and longitude from 180(E to 180(W is displayed in the right display. Resolution is 0.1
minute. The display is used when inserting present position during initialization of
the two IRUs or when monitoring present position from an IRU during flight.
WIND (Wind speed and Direction) - Wind speed from 0 to 256 knots is displayed in the
right display with a resolution of 1 knot. Wind direction form 0 to 359 degrees is
displayed in the left display with a resolution of 1 degree.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
A.
Display Selector Switch (DSPL SEL)
(Cont.)
HDG/STS (Heading/Status) - True heading from 0 to 359.9 degrees is displayed in the
left display with a resolution of 0.1 degree.
The ALIGN STATUS appears in the left side of the right display and is a count down of
the last 7 minutes of the ALIGN cycle.
Malfunction codes (M/C) appear in the right side of the right display.
TEST - The test position provides a remote test signal to the selected IRU.
BRT - The brightness control knob is concentric with the DSPL SEL switch and is a
potentiometer to control brightness of the displays.
B.
System Display Switch (SYS DSPL)
The system display switch selects the left or right IRU for display of navigational
data on the display.
C.
Keyboard
The twelve-key keyboard allows entry of initial latitude and longitude when in ALIGN
and of set-magnetic-heading when in ATT. The keyboard has 12 panel lamps for keyboard
lighting, which use the airplane light dimming control circuits. The ENT and CLR keys
have highlighter bars which function as follows:
When the N, S, E, W or H key is pressed, the ENT highlighter bar illuminates and
stays lit while digits are keyed in. When the ENT key is pressed, the ENT
highlighter bar extinguishes and a reasonableness check is done on the data. When
reasonable, the data is transmitted to the IRU and the display shows the selected
parameters. If the data is unreasonable, the CLR highlighter bar illuminates and
the display retains the unreasonable entry. Pressing the CLR key causes the CLR
highlighter bar to extinguish and the ISDU to display the selected parameters.
6. Monitor
The ISDU displays malfunction codes when abnormal condition are sensed by an IRU.
Malfunction codes are displayed in the right display when the DSPL SEL switch is placed to
HDG/STS.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
7. Self-Test
When held in the spring loaded TEST position, the DSPL SEL switch causes fixed outputs from
the IRU selected by the SYS DSPL switch. The outputs are displayed on the ISDU and the
associated pilot's instruments. The test is inhibited during the navigational mode, when
the airplane ground speed is greater than twenty knots, and during the attitude mode. The
test results in displays of values from memory for two seconds, fault messages for eight
seconds, followed by test value displays.
CAUTION:
FAILURE TO OBSERVE ELECTROSTATIC DISCHARGE PRECAUTIONARY PROCEDURES WHEN HANDLING
THE ISDU MAY RESULT IN DEGRADATION OR FAILURE OF THE ISDU.
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MAINTENANCE TRAINING MANUAL
INERTIAL SYSTEM DISPLAY UNIT
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
FMCS CONTROL DISPLAY UNIT
1. Purpose
The inertial reference system may be initialized by using either of the two flight
management computer system control display units (FMCS CDU). Initializing (present position
insertion) is accomplished when the inertial reference system is in alignment. In addition,
BITE procedures for the IRS are accomplished by using either CDU.
2. Location
The two FMCS CDUs are located on the forward electronic panel, P9.
3.
Physical Characteristics
Refer to the FMCS section for all other information on the CDU.
FMCS CONTROL DISPLAY UNIT
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MAINTENANCE TRAINING MANUAL
OPERATION
IRS INITIALIZATION
1. Operation
The inertial reference system is initialized by entering present position on the control
display unit of the flight management computer system (FMCS CDU) or by using the ISDU.
Initializing with the FMCS CDU or the ISDU requires the data be entered only once for the
IRUs currently in the alignment procedure. Additional entries during the ALIGN period are
optional. The IRUs use the latest entry for aligning.
2. Normal Sequence
A.
Initialization Procedure (FMCS CDU)
The IRUs are placed in ALIGN or NAV modes using the mode select switches on the MSU.
The ON DC annunciator and then the ALIGN annunciator illuminate for both IRUs.
When the airplane is on the ground and the IRS has not been initialized, pressing the
INIT/REF key displays the position initialization page. Other ways of accessing the
position initialization page are discussed in the FMCS portion of this course.
One of three ways is used to enter present position into line 4R of the CDU (SET IRS
POS line):
The latitude and longitude are transferred from the LAST POS line (1R) to SET IRS
POS through the scratch pad. (1R, then 4R are line selected).
REF AIRPORT (2L) data is entered and the resulting position data is transferred to
the SET IRS POS line through the scratch pad. (2R, then 4R are line selected).
The latitude and longitude is entered into the scratch pad using the alpha-numeric
keyboard. The scratch pad contents are then line selected into the SET IRS POS
line (4R).
When the IRSs have accepted the initialization latitude/longitude, the display on the ISDU
is monitored with PPOS selected on the DSPL SEL switch.
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MAINTENANCE TRAINING MANUAL
OPERATION
B.
Initialization Procedure (ISDU)
The DSPL SEL switch is placed in PPOS position.
The latitude and longitude of present position is entered with the keyboard. Either
latitude or longitude may be entered first.
For latitude, N2 or S8 key is pressed. The letter N or S appears on the left digit and
an 0 appears in the right digit of the left display window. The highlight bar on the
ENT Key illuminates.
Latitude entry is continued. As a key is pressed, the digit appears in the right digit
of the left display and the digits already entered shift one space to the left. ENT is
pressed to insert the latitude into the IRU computer. The highlight bar on the ENT key
goes out.
Longitude is entered in the right display in a similar way, starting with the W4 or E6
key. ENT is pressed to insert the display information into the IRU computer. The IRU
selected by the SYS DSPL switch should return the entered latitude and longitude to the
display.
If a latitude of greater than 90 degrees or a longitude of greater than 180 degrees or
a minute value of greater than 59.9 is used pressing the ENT key causes the highlight
bar on the CLR key to illuminate. Pressing the CLR key blanks the display and enables
another selection.
WARNING:
THE AIRPLANE MUST NOT BE MOVED WHEN THE IRU'S ARE IN THE ALIGN MODE.
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MAINTENANCE TRAINING MANUAL
IRS INITIALIZATION
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MAINTENANCE TRAINING MANUAL
IRS - REFERENCE AIRPORT PRESENT POSITION ENTRY
21
MAINTENANCE TRAINING MANUAL
NOTES
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
INERTIAL REFERENCE UNIT
1. Purpose
The inertial reference unit (IRU) provides attitude, acceleration, angular rates, velocity,
true and magnetic headings, positional data, absolute altitude and wind data signals. The
signals are developed from a set of three laser gyros and three accelerometers mounted to
airplane reference. The signals are provided to other systems including the flight
management computer system, the digital flight control system, the autothrottle, the VHF
navigation system and the pilots' instruments.
2. Location
The IRUs are located in the E & E compartment mounted on the E3-5 shelf.
3. Physical Description
The unit conforms to ARINC 600 for cooling and connections, is a 10 MCU size box and weighs
44 pounds (20 kgms). Internally, the module contains thirteen card assemblies, a power
supply, a sensor assembly containing the laser gyros and the accelerometers, and a wired
chassis assembly.
A pushbutton test switch, a fault ball indicator and an elapsed time indicator are mounted
on the front face of the unit.
The total time indicator displays accumulated time.
The IRU is interfaced with airplane wiring through a triple section connector of which only
the two lower sections are used.
Two hold down hooks are used to secure the IRU on its mount.
4. Power
The IRU has two power sources, one a 115 volt ac source and one a 28 volt dc source. Either
source is sufficient for operation.
5. Monitor
The fault ball indicator is controlled by internal monitoring circuits. A valid condition
is indicated by black and an abnormal condition by yellow. When the abnormal condition is
temporary, the indicator is reset to black by a power recycle.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
7. Self Test
When pressed, the test switch causes simulated outputs that are displayed on the ISDU and
the pilot's instruments. The test switch is inhibited during the navigational mode when the
airplane ground speed is greater than twenty knots and during the attitude mode. The test
results in displays of values from memory for two seconds, fault messages for eight seconds,
followed by test value displays.
CAUTION:
THE UNIT CONTAINS STATIC SENSITIVE DEVICES, AND PRECAUTIONS MUST BE USED WHEN
HANDLING THE UNIT. FAILURE TO OBSERVE ELECTROSTATIC DISCHARGE PRECAUTIONARY
PROCEDURES WHEN HANDLING THE IRU MAY RESULT IN DEGRADATION OR FAILURE OF THE IRU.
THE UNIT CONTAINS HIGH VOLTAGE DEVICES. THE UNIT COVER SHOULD NOT BE REMOVED
EXCEPT IN THE OVERHAUL SHOP.
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MAINTENANCE TRAINING MANUAL
INERTIAL REFERENCE UNIT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
DIGITAL ANALOG ADAPTER
1. Purpose
The Honeywell digital to analog adapter (DAA) is a component of the inertial reference
system (IRS) but is interfaced with other airplane systems for processing signals
independent of the status of the IRS. The two DAAs accept inputs in various formats,
primarily ARINC 429, and convert them to outputs compatible with various airplane systems
and instruments. The DAAs accept ARINC 429 inputs from the IRUs and convert the 429 format
into analog pitch, roll and magnetic heading signals with associated valids to control the
presentations on the pilots' ADIs and RDDMIs. The DAAs also interface the FMC and the VHF
navigation systems.
2. Location
The two DAAs are installed on the E1 rack, 2nd shelf in the electrical and electronic
compartment.
3. Physical Description
The DAA is a 1/2 ATR box with ARINC 404A specification for construction and cooling and
weighs 23.5 pounds (10.66 kgms). Two hold down hooks are used to secure the DAA in its
mount. There are no external controls or indicators on the front face. Internally, the
module contains 13 card assemblies including two cards supplying unit power. A dual, rear
mounted connector interfaces the unit with airplane wiring.
4. Power
Two 115 volt ac and two 28 volt dc power sources are required to power the DAA circuits.
5. Monitor
Abnormal operation of the DAA signal processing, such as pitch, roll and magnetic heading
causes a malfunction code display on the ISDU and a flag to appear on the associated
instrument.
6. BITE
The DAAs are checked during IRS BITE. The status of the DAAs is displayed during FMCS BITE.
CAUTION:
THE UNIT CONTAINS STATIC SENSITIVE DEVICES. OBSERVE PRECAUTIONS FOR HANDLING
ELECTROSTATIC SENSITIVE DEVICES. FAILURE TO OBSERVE ELECTROSTATIC DISCHARGE
PRECAUTIONARY PROCEDURES WHEN HANDLING MAY RESULT IN DEGRADATION OR FAILURE OF THE
DAA.
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DIGITAL ANALOG ADAPTER
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MAINTENANCE TRAINING MANUAL
OPERATION
IRU DISPLAY DATA - ADI
Operation
Power
The 26 volt ac power is utilized to power the ADI internal power supply and furnish a
reference voltage for conversion of data.
Attitude Operation
Pitch and roll attitudes are computed parameters transmitted by the IRU in a digital
format. The DAA transforms the digital attitude data to an analog (syncro) format and
sends it to its onside ADI and alternate attitude data to the offside attitude transfer
relay. The transformation of the primary and alternate attitude data is accomplished in
two totally independent internal functions of the DAA (ESSENTIAL/NONESSENTIAL);
accompanying each attitude parameter is a VALID generated by a BITE function in
conjunction with a monitor channel to assure the integrity of each output.
The ADI includes two servocircuits that control the horizon and a flag circuit that
monitors the status of the magnet heading source and the operation of the
servocircuits. Three wire attitude information is applied to synchro-resolvers which
each supply an output to a demodulator and amplifier which drives the motors and
repositions the synchroresolver rotors. The motors also position the rotors of two
synchro transmitters that are used in some configurations to supply attitude
information to the flight data acquisition unit. An AND gate, which controls the ATT
flag, monitors excitation from the synchroresolvers, proper follow-up on commands by
the servocircuits, dc from the power supply and the valid from the attitude source.
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MAINTENANCE TRAINING MANUAL
IRU DISPLAY DATA – ADI
29
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
IRU DISPLAY DATA - HSI II
Operation
Power
26 volt ac power is utilized to enable the internal power supplies and to furnish
reference voltages for conversion data.
Data Source Selection
The data sources for drift angle and ground speed are the IRU in VOR/IRS mode and the
FMC in NAV mode. Desired track is supplied by the FMC. All three parameters are
received by the DAA. Source selection is accomplished in the DAA by means of an input
from the radio/nav relay.
The desired track output from the DAA is displayed in NAV mode. In VOR/ILS mode,
selected course from the AFCS mode control panel is displayed. This input selection
into the HSI also is accomplished by means of the respective radio/nav relay.
DAA Operation
The DAA transforms the received digital DA and desired-track data to analog (synchro)
format, and received digital ground-speed to ARINC 561 six-wire digital format. The
transformation of the data is accomplished in the NON-ESSENTIAL internal function of
the DAA.
The BITE function generates a NAV VALID used to assure groundspeed data integrity, and
a DA VALID to assure drift-angle data integrity.
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MAINTENANCE TRAINING MANUAL
OPERATION
HSI Operation
The HSI includes servocircuits that control the drift-angle and the desired-track
indicators, and a flag circuit that monitors the status of the A source and the operation
of the servocircuits. Three-wire data information is applied to the respective synchroresolvers which supply outputs to demodulators and amplifiers which drive the indicator
drive motors and reposition the synchro-resolver rotors. A DA VALID maintains the DA
input circuit contacts closed; loss of DA VALID opens these contacts and causes the
servocircuit to drive the drift-angle indicator to the 6 o'clock position behind the
mask.
Ground speed information is received on an ARINC 561 six-wire digital bus and displayed on
an LED display. Loss of the NAV VALID signal causes the display to blank. If groundspeed
is NCD, bit 32 of the data word is set which causes dashes to be displayed.
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33
IRU DISPLAY DATA – HSI II
MAINTENANCE TRAINING MANUAL
OPERATION
IRU DISPLAY DATA – RDDMI
Operation
Power
The 26 volt ac power is utilized to enable the RDDMI internal power supply and furnish
a reference voltage for conversion of data.
Heading Operation
Heading is a computed parameter transmitted by the IRU in a digital format. The DAA
transforms the digital heading data to an analog (syncro) format and sends it to its
onside RDDMI and alternate heading data to the offside IRS transfer relay. The
transformation of the primary and alternate heading data is accomplished in two totally
independent internal functions of the DAA (ESSENTIAL/NONESSENTIAL) accompanying each
heading parameter is a VALID generated by a BITE function in conjunction with a monitor
channel to assure the integrity of each output.
The RDDMI includes a servocircuit that controls a magnetic heading card and a flag
circuit that monitors the status of the magnet heading source and the operation of the
servocircuit. Three wire heading information is applied to a synchro-resolver which
supplies an output to a demodulator and amplifier which drives the motor and card and
repositions the synchroresolver rotor. The motor also positions the rotor of a synchro
transmitter that is used in some configurations to supply magnetic heading to a
navigational system. An AND gate, which controls the HDG flag, monitors excitation
from the synchroresolver, proper follow-up on commands by the servocircuit, dc from the
power supply and the valid from the heading source.
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MAINTENANCE TRAINING MANUAL
IRU DISPLAY DATA – RDDMI
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NOTES
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MAINTENANCE TRAINING MANUAL
Vertical Speed Indicator Schematic
EFFECTIVITY
AIRPLANES WITH INERTIAL
VERTICAL SPEED INDICATOR
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
IRS MASTER CAUTION UNIT AND LIGHT
1. Purpose
The IRS Master Caution Unit senses for Inertial Reference System (IRS) abnormal conditions
and controls the IRS master caution light on the left side of the glareshield. The unit
also supplies an input to the annunciator and dim module to operate the left and right
master caution lights.
2. Location
The IRS Master Caution Unit is installed on the P10 panel on the right side of the flight
control cabin.
3. Physical Description
The unit contains four cards, one card each for monitoring ON DC, FAULT and DC FAIL logics,
from each IRU and one output card for control of the IRS caution light.
4. Test
Part of the IRS master caution unit and the IRS lamp on the left master caution annunciator
may be tested by the master dim test circuit.
IRS MASTER CAUTION UNIT AND ANNUNCIATOR
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
FLIGHT INSTRUMENT ACCESSORY UNIT
1. Purpose
The flight instrument accessory unit contains auxiliary circuits that complete the circuits
of a number of navigational systems to the configuration of the airplane.
2. Location
The unit is installed on the E1-2 shelf in the E&E compartment.
3. Physical Description
The unit contains relays, resistors, diodes, circuit boards and other components that
integrate several navigation systems to the airplane. For the inertial reference system,
the unit contains a relay circuit that controls the DC applied to the ground crew call horn.
The circuit activates the horn after a 19-second delay should the IRS be placed on battery
power with the airplane on the ground. For IRU 2, a second relay circuit limits DC power to
IRU 2 for five minutes when there is no ac power to IRU 2.
FLIGHT INSTRUMENT ACCESSORY UNIT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
ATTITUDE DIRECTOR INDICATOR (PITCH AND ROLL INDICATIONS)
1. Purpose
The attitude director indicator (ADI) is a multipurpose display instrument which displays
pitch and roll attitudes to the pilots during the entire flight.
2. Location
One ADI is installed in the center of each pilots' instrument panels, P1 and P3, in the
flight control cabin.
3. Physical Characteristics
The ADI conforms to ARINC 408, is a five inch instrument and weighs approximately nine
pounds. Pitch indications range from 90 degrees nose up to 90 degrees down. Roll
indications range around 360 degrees.
Two rear mounted connectors provide an interface of the ADI with airplane wiring.
A test switch on the lower right of the instrument is used to test the attitude indications.
4. Power
The pitch and roll servocircuits require 26 volt ac to operate.
5. Monitor
The ATT flag, solenoid controlled, is black letters on a red background and monitors the ADI
power source, the pitch and roll servocircuits and the attitude source. In addition, the
two ADIs are monitored for pitch and roll comparison by the instrument comparator.
6. Self Test
Pushing the TEST switch results in the indication of 10 degrees pitch up and 20 degrees
right roll with ATT flag in view.
CAUTION:
FAILURE TO OBSERVE ELECTROSTATIC DISCHARGE PRECAUTIONARY PROCEDURES WHEN HANDLING
THE ADI MAY RESULT IN DEGRADATION OR FAILURE OF THE ADI.
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MAINTENANCE TRAINING MANUAL
ATTITUDE DIRECTOR INDICATOR (PITCH AND ROLL INDICATOR)
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
HORIZONTAL SITUATION INDICATOR (IRS INPUT)
1. Purpose
The horizontal situation indicator (HSI) provides the pilots with a display of navigational
data from a selected inertial reference unit (IRU).
2. Location
One HSI is located on each of the pilots' instrument panels, P1 and P3.
3. Physical Description/Features
The HSIs are five-inch analog-type instruments.
the following indicators:
They display IRU navigational data with
Magnetic Heading
Ground Speed Readout
Drift Angle Pointer
A.
Magnetic Heading
Magnetic heading is continually controlled from the IRU and is read against the lubber
line through 360 degrees in five degree increments.
B.
Ground Speed Readout
The ground speed readout indicates ground speed computed by the IRU when the pilot
selects VOR/ILS reference for the HSI.
C.
Drift Angle Pointer
The drift angle pointer indicates IRU drift angle when the pilot selects VOR/ILS
reference for the HSI. Drift angle is read relative to the lubber line.
4. Control
Selection of VOR/ILS or NAV as the signal source for navigation data is controlled by the
HSI transfer switches. The HSI transfer switches are located on the glareshield.
In the VOR/ILS position, the VHF navigation receiver and the IRU supply the signals. In the
NAV position, the FMCS supplies the signals.
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COMPONENT FUNCTIONAL DESCRIPTION
5. Input
The data input to the HSI from the IRU is converted to analog in the DAA. A discrete is
received by the DAA from the HSI transfer switch for navigation source selection.
6. Monitor
The HSI monitors its various indicating circuits with flags. Several of the flags are also
be set by external source failure.
A.
Magnetic Heading
The HDG flag, black letters on red, monitors the operation of the HSI processing
circuits for heading display and the selected IRU source.
B.
Ground Speed
Failed ground speed signal source or internal circuits causes the readout to indicate
dashes.
C.
Drift Angle
Failure of the drift angle source causes the drift angle indicator to stow behind the
drift angle mask.
D.
VOR/ILS-NAV Annunciator
The VOR/ILS-NAV annunciator indicates VOR/ILS for radio/compass and IRU signal sources
and NAV for FMC signal sources.
E.
MAG/TRUE Annunciator
The MAG/TRUE annunciator indicates which heading is displayed from the IRU. The
annunciator always shows MAG when no MAG/TRUE switch is installed in the airplane.
CAUTION:
FAILURE TO OBSERVE ELECTROSTATIC DISCHARGE PRECAUTIONARY PROCEDURES WHEN HANDLING
THE HSI MAY RESULT IN DEGRADATION OR FAILURE OF THE HSI.
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HORIZONTAL SITUATION INDICATOR (IRS) – INPUTS
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COMPONENT FUNCTIONAL DESCRIPTION
RADIO DIGITAL DISTANCE MAGNETIC INDICATOR (IRS INPUTS)
1. Purpose
The radio digital distance magnetic indicator (RDDMI) is a multipurpose instrument which
displays magnetic heading from a selected IRU.
2. Location
One RDDMI is installed on each of the pilots' instrument panels, P1 and P3.
3. Physical Description
The instrument contains a servodriven card that displays magnetic heading through 360
degrees under a lubber line, in five degree increments, and a HDG flag, black letters on
red. The RDDMI weighs 5.5 pounds (2.5 kgs) and has three rear mounted connectors.
4. Power
26 volts ac is required to power the servomechanism that drives the compass card.
5. Monitor
The HDG flag monitors the operation of the compass card drive and the magnetic heading
source.
CAUTION:
FAILURE TO OBSERVE ELECTROSTATIC DISCHARGE PRECAUTIONARY PROCEDURES WHEN
HANDLING THE RDDMI MAY RESULT IN DEGRADATION OR FAILURE OF THE RDDMI.
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MAINTENANCE TRAINING MANUAL
RADIO DIGITAL DISTANCE MAGNETIC INDICATOR
(IRS INPUTS)
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COMPONENT FUNCTIONAL DESCRIPTION
ATTITUDE AND COMPASS TRANSFER SWITCHES
1. Purpose
The attitude and compass transfer switches enable either pilot to transfer his instruments
to an alternate navigation system source. The switches also provide discrete inputs to the
digital flight control system (DFCS) to provide attitude and heading input selection within
the DFCS.
2. Location
The switches are installed on the instrument switching panel, located on the forward
overhead panel P5.
3. Physical Description
The switches are three position switches, fixed at each position. The three positions are
NORMAL, BOTH ON 1 and BOTH ON 2.
4. Control
The attitude transfer switch controls two relays, R51 and R67, the attitude transfer relays,
which are located in the E&E compartment mounted on rack 1, second shelf. The compass
transfer switch controls two relays, R14 and R15, the compass transfer relays, located in
the same area as the attitude transfer relays.
5. Monitor
Operation of the switches and the relays is monitored by the flags in the pilots'
instruments.
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ATTITUDE AND COMPASS TRANSFER SWITCHES
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OPERATION
IRS GENERAL THEORY
Operation
The IRS provides basic heading and attitude reference information by utilizing computations
based on accelerometer and laser gyro sensed signals. Three accelerometers and three laser
gyros are used. The accelerometers are positioned in the inertial reference units so that
they are oriented along the x, y and z axes of the airplane. This orientation allows the
IRU to sense accelerations along each of the three axes.
The three laser gyros are positioned to sense pitch, roll and yaw rotation around the x, y,
and z axes.
These sensors are in the strap-down configuration. That is, they are oriented with respect
to each of the aircraft's 3 axes and are not allowed to move when the aircraft rotates about
or accelerates along any of these axes.
Computer manipulation of the signals from all six sensors provide the basic heading and
attitude reference signals along with present position, accelerations, ground speed, drift
angle and attitude rate information.
The first requirement which must be met for proper IRS operation is alignment. IRS
alignment basically consists of determination of local vertical and true north.
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SENSOR ASSEMBLY
• INERTIAL SENSORS FIXED
RELATIVE TO THE STRUCTURE
• RATE OUTPUTS
• ACCELERATION OUTPUTS
IRS GENERAL THEORY
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OPERATION
IRS THEORY OF OPERATION - ALIGNMENT
Operation
IRS alignment consists of determining local vertical and initial heading. Both accelerometer
and laser gyro inputs are used for alignment. The alignment computations use the basic
premise that the only accelerations during alignment are caused by the earth's gravity. The
only motion during alignment is caused by the earth's rotation. Accelerations because of
gravity are always perpendicular to the earth's surface and define the local vertical. The
local vertical is used to compute the attitude data so that it is accurately referenced to
vertical. Initially, only a coarse vertical is established. Once vertical is established,
the laser gyro sensed earth rate components are used to establish the heading of the
airplane. As the alignment continues, both the vertical reference and the heading
determinations are fine tuned for maximum accuracy. The vertical axis orientation for
attitude reference, relative to the earth's surface, is based on airplane position input to
the IRU. The initial position entry be should made any time during the alignment period.
Earth rate sensing by the laser gyros allows the IRU to determine initial latitude. The
gyro determined latitude is compared to the crew entered latitude. Crew entered longitude
is compared to the last stored longitude. The comparisons must be within limit to complete
the alignment period, during which all outputs of the IRU, except for present position, are
set to NCD (no computed data). The minimum duration of the alignment mode is 10 minutes.
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IRS THEORY OF OPERATION - ALIGNMENT
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OPERATION
IRS OPERATION THEORY OF OPERATION - NAVIGATION MODE
Operation
In the navigate mode the IRUs provide outputs for attitude, heading, accelerations, ground
speed, position and drift angle. These outputs are all derived from gyro and accelerometer
strap down sensor data. The initial attitude, heading and velocity signals are modified by
inputs from the sensors to establish real time present parameters through integration and
computer calculations. Additional calculations by the computer establish such parameters
as present position, ground speed and drift angle. Inputs from the ADCs are used for
inertially smoothed altitude and altitude rate (baro altitude) and wind speed/direction
(true airspeed).
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IRS THEORY OF OPERATION - NAVIGATE MODE
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OPERATION
IRS - SINGLE AXIS NAVIGATION COMPUTATIONS
Operation
One of the types of sensing devices within the IRU is an accelerometer. The accelerometer
is shown as a mass centered in a case by two springs. As the airplane accelerates, the
mass is displaced from center, causing an electrical pickoff signal to be generated. This
signal is amplified and applied as feedback to recenter the mass. The amount of signal
required to keep the mass centered is therefore proportional to acceleration. This
recentering operation allows the accelerometer to sense over a wide range and also be able
to sense very small changes in acceleration.
The recentering signal is integrated once to give velocity and integrated a second time to
give distance.
The airplane's present position is calculated by adding the distance flown to the starting
position.
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Accelerometer
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OPERATION
LASER GYRO
Operation
Laser Beam Generation
Each gyro is a triangular-shaped, helium-neon laser that produces two light beams, one
traveling in the clockwise direction and one in the counterclockwise direction.
Production of the light beams, or lasing, occurs in the gas discharge region by
ionizing the low pressure mixture of helium-neon gas with high voltage to produce a
glow discharge. Light produced from the lasing is reflected around the triangle by
mirrors at each corner of the triangle to produce the clockwise and counterclockwise
light beams. One of the three corners contains the corner prism to allow the two beams
to mix together to form the interference fringe pattern on the readout detector.
Angular Rate Detection
As long as the gyro is stationary, the fringe pattern is also stationary since the
frequencies of both light beams are the same. When the gyro rotates about the axis
perpendicular to the lasing plane, the two light beam frequencies are slightly
different during the rotation. The two frequencies are different because the two paths
traveled by the beams are different. This difference in frequencies causes the fringe
pattern to move left or right depending on which direction the gyro rotates. Photo
diodes in the readout detector convert the fringe pattern movement into electrical
pulses corresponding to the rotation rate.
Dither Motor
At low rotation rates, the two light beams get coupled together in a condition called
laser lock-in. To avoid loss of information at these low rates, a piezo-electric
dither motor vibrates the gyro assembly through the lock-in region. The motions caused
by the dither motor are optically decoupled from the gyro output. The dither motor
vibrations can be felt on the IRU case and produces an audible hum.
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LASER GYRO
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NOTES
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OPERATION
IRS BLOCK DIAGRAM
Operation
The IRS block diagram shows the interface of the components of the IRS and between the IRUs
and other airplane systems.
Power
115 volts ac and 28 volts dc are available to each IRU for normal operation. Each DAA
requires two 115 volt and two 28 volt sources for operation.
Mode Selection
The MSU supplies mode selection into the IRUs which energizes the IRU power supplies
and arms mode progression.
Bus Information
Communication between the IRS components and with other systems is accomplished through
ARINC 429 busses.
ISDU Outputs
The ISDU supplies operator inserted latitude and longitude to the IRUs during alignment
and magnetic heading during attitude mode. It also supplies a test discrete for system
self-test in parallel with an FMC initiated test through the DAAs.
ADC Outputs
The ADCs supply altitude, altitude rate and true airspeed to the IRUs.
FMC Outputs
The FMC bus outputs to the IRUs contain operator inserted latitude and longitude during
alignment and magnetic heading during attitude mode. In addition, a BITE word is made
available to the IRUs during BITE through the DAAs.
IRU Outputs
The IRUs each supply three busses to interfacing equipment. IRU-1 bus is supplied to
the DAA. The DAA reformats ground speed, magnetic heading, pitch, roll and yaw rate to
analog signals for use in other systems. IRU-2 bus is supplied to the flight control
computers (FCC) which extract discretes and many parameters, including pitch and roll.
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OPERATION
Operation
IRU Outputs (Cont.)
The IRU-3 bus is applied to the DAAs, to the FCCs as an alternate to the IRU-2 bus, to
the ISDU which extracts information for display of navigational parameters, to the FMC,
TO THE A/T computer and to the weather radar R/T unit.
DAA Outputs
The DAAs convert digital words of pitch, roll and magnetic heading into three wire
analog signals and apply them through the transfer relays, pitch and roll to the ADIs,
and magnetic heading to the RDDMIs for display. The DAAs also convert navigational
information from the IRU and the FMC to a BCD output for display on the HSI. Roll
attitude from DAA 1 is applied to the Y/D coupler for turn coordination.
Magnetic heading references from the DAAs are supplied to the DFCS mode control panel
and applied to the heading selector and the course selectors for control of the
indicators in the pilots' HSIs.
The VHF NAV receivers are normally supplied with magnetic heading reference from the
DDAs. The alternate sources are the pilots' RDDMIs.
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OPERATION
NORMAL IRS ALIGNMENT PROCEDURE
Operation
General
During the alignment process, the IRU determines the local vertical and the direction
of true north.
CAUTION: THE AIRPLANE MUST NOT BE MOVED DURING ALIGNMENT.
Normal Procedure
Alignment is achieved by the procedure shown. Normally alignment takes ten minutes,
after which the IRUs enter into the NAV mode. The operator must insert present
position sometime during the alignment process, using either the FMCS CDU or the ISDU.
Problems with the alignment process are indicated by a flashing ALIGN annunciator or
steady FAULT annunciator on the MSU.
The IRU automatically advances to the navigation mode at completion of the ten minute
alignment provided the present position has been entered and the NAV mode is selected.
If present position has not been entered by the time alignment is complete, the ALIGN
annunciator flashes.
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NORMAL IRS ALIGNMENT
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OPERATION
IRS ALIGNMENT INDICATIONS - EXCESSIVE MOTION
Operation
When the airplane is moved while the IRUs are in ALIGN, a status code 3 is displayed on the
ISDU and "IRS MOTION" is displayed in the CDU scratchpad. Thirty seconds after the airplane
motion has stopped the IRUs automatically restart another alignment. Fault code 3 will then
go blank. The "IRS MOTION" message will blank. The new alignment will then take 8 minutes
to complete.
IRS ALIGNMENT INDICATIONS - EXCESSIVE MOTION
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OPERATION
IRS ALIGNMENT INDICATIONS - LAT/LON ONE DEGREE DIFFERENCE –
(During Alignment Period)
Operation
If the entered latitude or longitude differs by more than 1º from the IRU stored value, the
ALIGN light will begin flashing. With the DSPL SEL switch in the HDG/STS position, status
code 4 (align fault) will be displayed. Entering the same LAT/LON again will cause the
ALIGN annunciators to illuminate steadily and status code 4 will disappear.
When the IRU has been removed from the airplane and serviced in an avionics repair facility,
no comparison is done on entered LAT/LON and the IRU stored value for LAT/LON. This
prevents the situation described above when an IRU is removed from the airplane for repairs
and reinstalled in the airplane at a different location.
IRS ALIGNMENT INDICATION - LAT/LON 1º DIFFERENCE
(DURING ALIGNMENT) PERIOD
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OPERATION
IRS ALIGNMENT INDICATIONS - LATITUDE DISAGREEMENT - (AT THE END OF THE ALIGNMENT PERIOD)
Operation
When the entered latitude does not agree with the latitude computed at the end of the tenminute alignment period, the ALIGN light on the MSU flashes indicating operator attention.
ENTER IRS POS is displayed on the CDU/ANCDU. Enter the correct position. If the latitude
disagreement persists, the ALIGN and FAULT lights illuminate steadily and status code "02"
is displayed. CYCLE IRS OFF-NAV is displayed on the CDU/ANCDU. Turn the IRU off. After
the ALIGN light extinguishes, turn the IRU on and enter known accurate present position
(twice, if required, using indentical entries). If the status code "02" persists, replace
the IRU.
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IRS ALIGNMENT INDICATIONS - LATITUDE DISAGREEMENT
(AT THE END OF THE ALIGNEMENT PERIOD)
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NOTES
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OPERATION
IRS ALIGNMENT INDICATIONS - NO PPOS ENTERED
Operation
When a present position has
annunciator flashes, status
on the CDU/ANCDU. Once the
extinguish and the NAV mode
not been entered by the end of alignment period, the ALIGN
code 08 is displayed on the ISDU, and ENTER IRS POS is displayed
present position is entered, the ALIGN annunciator will
is entered.
IRS ALIGNMENT INDICATIONS - NO PPOS ENTERED
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OPERATION
IRS ATTITUDE MODE
Operation
General
During ATT mode, the IRU functions as a combined directional gyro and vertical gyro.
The attitude mode is entered by moving the mode select switch to ATT. The mode is used
only where attitude information is needed, or when the navigation function of the IRU
fails during flight but the attitude functions remains operational. This condition
occurs when the IRU loses all power during flight. The ATT mode is the only position
of the mode select switch that the IRU accepts set magnetic heading input from the ISDU
or FMCS CDU. When the mode select switch is moved directly from OFF to ATT, attitude
output is available immediately with full performance available within 30 seconds.
Once ATT mode is selected, the IRU is latched into attitude mode until power is removed
by selecting the OFF position.
Heading Outputs
When ATT is selected and a magnetic heading output is desired, the heading is
initialized through the FMCS CDU or the ISDU. With magnetic heading initialized, the
IRU uses the initial magnetic heading output and changes magnetic heading output as the
platform heading changes. The code 09 appears in the right display of the ISDU until
the magnetic heading is initialized.
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IRS ATTITUDE MODE
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MAINTENANCE PRACTICES
IRS STATUS/MALFUNCTION CODES
Fault Isolation
System component status/malfunctions are indicated on the right display when the display
select switch is in HDG/STS. The table shows a code number indicating a particular
component status, or for condition. ISDU power loss, number 10, is a normal indication
during initial power on when one IRU is powered and the other is not.
The component associated with IRU FAIL or ADC FAULT is a function of the SYS DSPL switch
position.
If multiple codes exist, pressing the CLR key will enable all to be viewed in sequence until
the original code reappears.
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01 ISDU FAIL
02 IRU FAIL
03 EXCESSIVE MOTION IN ALIGN MODE
04 ALIGN FAULT
05 DAA FAULT L
06 DAA FAULT R
07 ADC FAULT
08 ENTER PRESENT POSITION
09 ENTER HEADING
10 ISDU POWER LOSS
IRS STATUS/ MALFUNCTION CODES
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MAINTENANCE PRACTICES
IRU REJECTION CRITERIA
Position Error
Position error is the distance between the airplane's actual position (typically
latitude/longitude of an assigned gate). If the position error at the end of two
consecutive flights is greater than (3 + 3T) nm, T = time in NAV mode, the IRU should be
removed.
Residual Ground Speed Error
Residual ground speed is the ground speed calculated by the IRU after a flight while the
airplane is stationary. This ground speed is displayed on the FMCS CDU/ANCDU on page two of
POS INIT. If this ground speed is greater than 20 knots for two consecutive flights, the
IRU should be removed.
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IRU REJECTION - CRITERIA
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NOTES
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1. DETERMINE IRS RADIAL POSITION ERROR
NOTE:
1. TO ACCOMPLISH THE FOLLOWING STEPS, EACH IRS MODE SELECT SWITCH MUST REMAIN IN THE NAV POSITION AT FLIGHT
TERMINATION. THIS IS TO PRESERVE THE IRS POSITION COORDINATES COMPUTED DURING FLIGHT. THESE COORDINATES ARE
DISPLAYED ON THE POS REF 2/2 PAGE.
2. IN THE FOLLOWING STEPS AIRPLANE PRESENT POSITION AND IRS POSITION ARE ENTERED INTO THE FMCS ACT RTE LEGS 1/3
PAGE. WHEN ENTERED, THESE POSITIONS ARE DISPLAYED AS WAYPOINTS.
A.
PRESS FMCS CDU LEGS KEY. OBSERVE ACT RTE LEGS 1/3 PAGE IS DISPLAYED.
B.
DELETE ALL WAYPOINTS FROM ACT RTE LEGS 1/3 PAGE.
C.
ENTER AIRPLANE PRESENT POSITION COORDINATES INTO DATA FIELD OF ACT RTE LEGS 1/3 PAGE.
D.
PRESS CDU INIT REF KEY, THEN PRESS NEXT PAGE KEY. OBSERVE POS REF 2/2 PAGE IS DISPLAYED.
E.
TRANSFER IRS L POSITION COORDINATES, FROM POS REF 2/2 PAGE INTO THE DATA FIELD OF ACT RTE LEGS 1/3 PAGE.
F.
DISTANCE BETWEEN COORDINATES IS DISPLAYED IN NAUTICAL MILES, ON ACT RTE LEGS 1/3 PAGE. NOTE THE DISTANCE. THE DISTANCE
REPRESENTS RADIAL POSITION ERROR BETWEEN AIRPLANE PRESENT POSITION AND IRS L POSITION.
G.
DELETE IRS L POSITION FROM ACT RTE LEGS 1/3 PAGE.
H.
TRANSFER IRS R POSITION COORDINATES, FROM POS REF 2/2 PAGE INTO THE DATA FIELD OF ACT RTE LEGS 1/3 PAGE.
I.
DISTANCE BETWEEN COORDINATES IS DISPLAYED IN NAUTICAL MILES, ON ACT RTE LEGS 1/3 PAGE. NOTE THE DISTANCE. THE DISTANCE
REPRESENTS RADIAL POSITION ERROR BETWEEN AIRPLANE PRESENT POSITION AND IRS R POSITION.
2. REFER TO RADIAL POSITION ERROR CHART BELOW. AS A FUNCTION OF HOURS IN NAV MODE AND DISTANCES NOTED, DETERMINE IF RADIAL POSITION
ERROR IS EXCESSIVE OR WITHIN ACCEPTABLE LIMITS.
3. THE EQUATION 3+3T MAY ALSO BE USED TO DETERMINE IF RADIAL POSITION ERROR IS EXCESSIVE OR WITHIN ACCEPTABLE LIMITS. T IS HOURS
IN NAV MODE. DISTANCE ON ACT RTE LEGS 1/3 PAGE IS ACCEPTABLE IF LESS THAN 3+3T NAUTICAL MILES.
IRS Accuracy Criteria 1
EFFECTIVITY
AIRPLANES WITHOUT EFIS
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MAINTENANCE PRACTICES
IRS BITE INTRODUCTION
1. Tests
The IRS built-in test equipment (BITE) provides for the ground maintenance testing of the
IRS.
IRS BITE is contained in the DAA, and is accessed through the flight management computer
system CDU. In order to minimize ground test time, the BITE program is divided into three
fault isolation routines: current status, interface tests, and in-flight faults storage and
display.
2.
BITE Operation
See Chapter 22 BITE Section, of the Maintenance Manual, for complete details of the IRS BITE
operation.
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FMCS CDU - IRS Functions 1
EFFECTIVITY
AIRPLANES WITHOUT EFIS
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IRS BITE INTRODUCTION
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Basic Pipe Selection IRS BITE
Figure 10
EFFECTIVITY
AIRPLANES WITHOUT EFIS
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NOTES
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IRS DITF - In-Flight Faults and Current Faults
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IRS DITF - Interface Check L IRS
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IRS SELF TEST
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IRS DITF - Interface Check L - R IRS
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MAINTENANCE PRACTICES
IRS SELF-TEST
1. Test Preparation
An IRS self-test is performed from the ISDU or from the IRU. Before the test is performed,
certain conditions must be established. In addition to powering the associated systems, the
IRU to be tested must be in the ALIGN or NAV mode and the airplane must have a ground speed
of less than 20 knots.
2. Test
A.
Initiated at ISDU
With conditions for the self-test set, and the IRS selected with the SYS DSPL switch
the test is initiated from the ISDU by holding the DSPL SEL in the TEST position. For
the first two seconds, the following is observed on the ISDU and MSU:
All IRS DISPLAY Segments ON
ENT and CLR Cues ON
ALIGN, FAULT, ON DC and DC FAIL Annunciators ON
The ISDU and MSU displays and annunciators are OFF after the first two seconds. For
the next eight seconds (2-10 seconds after test start), the following indications are
observed on the instruments:
ATT flag in View on ADI
HDG flag in View on RDDMI and HSI
GND SPD flag in View and DA Pointer behind Mask on HSI when in
VOR/ILS Mmode for the HSI
After 10 seconds, the following values are observed on the instruments:
Pitch Attitude 5° UP
Roll Attitude 45° RIGHT
Magnetic Heading 15°
Drift Angle 10° LEFT
Ground Speed 200 knots
The test values exist as long as the DSPL SEL is held in the TEST position. The TEST
position of the DSPL SEL switch is spring-loaded and must be held in the TEST position.
When the DSPL SEL is released from TEST before 10 seconds, the test continues until 10
seconds after it was initiated and then stop.
When the test is ended, the instrument displays return to actual IRU values when the
IRU is in the NAV mode. When the IRU is in the ALIGN mode, the instruments will
continue to display the test values but flags are in view because the IRU output in
ALIGN is NCD (no computed data).
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MAINTENANCE PRACTICES
B.
Initiated at IRU
The self-test is initiated at the IRU by pressing and holding the test switch. The
test results for the first 10 seconds of the test is identical to the results from the
ISDU initiated test because the TEST switches are wired in parallel. After 10 seconds,
the instrument displays are the same as from the ISDU initiated test. In addition,
seven test values can be checked on the ISDU by rotating the SYS DSPL to the system
under test and rotating the DSPL SEL switch through its four positions. The values
which can be checked on the ISDU are:
True Track Angle 0.0°
Ground Speed 200 knots
Latitude N22° 30.0 min
Longitude E22° 30.0 min
True Wind Direction 30°
Wind Speed 100 knots
True Heading 10.0°
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EFIS Differences
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COMPONENT FUNCTIONAL DESCRIPTION
ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI)
Purpose
The EADI displays pitch and roll attitudes from the IRS when the IRUs are aligned and IRS
ground speed if the FMC fails.
Display data for each IRU condition is illustrated on the graphic.
Operation
In normal operation the display shows the attitude ball and reference scales for pitch and
roll. Ground speed from the IRU will be displayed if the ground speed from the FMC is
failed. With the airplane on the ground and the IRU's not initialized, the ground speed
display will be dashes. When the IRU's are initialized, the ground speed display will be
zero. With the airplane on the ground and the IRU's off, the ground speed display will be
blank. When attitude from the IRU is no computed data, the roll reference scale will be
displayed, and the attitude ball will be removed. If the attitude data from the IRU is
failed, the attitude ball is removed and the attitude flag is displayed. The roll reference
scale is also displayed.
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IRS DISPLAY DATA - EADI
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COMPONENT FUNCTIONAL DESCRIPTION
IRS DISPLAY DATA - EHSI/VOR-ILS
Operation
IRS generated display data is illustrated on the graphics for the normal, no-computed-data,
and failed conditions of the IRU. Airplane heading comes from the IRS, while wind
direction, wind speed and airplane track may come from the flight management computer (FMC)
or the IRS. As pointed out on the graphic, the FMC input data is assumed to be invalid for
the purpose of illustrating IRS inputs only.
The NAV mode is shown in the "FULL ROSE" format while the VOR mode is shown in the "EXPANDED
ROSE" format.
The EHSI displays the following IRS generated data:
Heading
Track Angle
Drift Angle
Wind (direction and magnitude)
Any other information shown is typical information that would be displayed but that does not
originate from the IRU.
The wind and track information is displayed from data out of the FMC (primary) or out of the
IRU (backup).
If the heading data from the IRU is No-Computed-Data then the heading display on the EHSI
will display dashes, the numbers on the compass rose are removed, and the track pointer is
removed.
If the heading data from the IRU is failed and a heading up mode is selected, then the
compass rose and all indicators related to heading are removed. The heading flag is
displayed at the top of the EHSI.
If heading data is failed and a track up mode is selected, the heading pointer will be
removed.
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IRS DISPLAY DATA - EHSI/VOR-ILS
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NOTES
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OPERATION
IRS BLOCK DIAGRAM
Operation
The IRS block diagram shows the interface between the various components of the Inertial
Reference System, and the interface between the IRS and other airplane systems.
Power
The IRU operates on 115 volts ac normally. However, if 115 volts is not available, the
IRU can operate on 28 volts dc from the battery. Switching occurs automatically in the
IRU.
Mode Selection
Mode select data from the MSU to the IRU enables the power supply in the IRU and
selects the operating mode of the IRU.
Bus Information
Communication between the various components of the Inertial Reference System and
between the IRS and other airplane systems is accomplished using ARINC 429 busses.
ISDU Outputs
The ISDU supplies entered present position to the IRUs during alignment and magnetic
heading during attitude mode. It also supplies a test discrete for system self-test.
Air Data Computer Outputs
The air data computers supply uncorrected altitude, altitude rate and true airspeed to
the IRUs.
FMC Outputs
The FMC bus outputs to the IRUs contain entered present position during alignment and
magnetic heading during attitude mode. The FMC also provides a digital BITE signal to
the DAA. The DAA decodes this signal and sends a test discrete to the IRU.
IRU Outputs
Each IRU has three output data busses supplying data to the interfacing equipment.
Each data bus contains the same information.
IRU bus No. 1 supplies data to the DAA, vertical speed indicator, and EFIS symbol
generator.
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OPERATION
IRU Outputs (Cont.)
IRU bus No. 2 supplies data to the Flight Control Computer.
IRU bus No. 3 supplies data to the DAAs, the ISDU, to the FMCs, the TCAS computer, and
to the stall management computers. IRU bus No. 3 also provides data to the offside
VSI, FCC, and EFIS symbol generator. IRU bus No. 3 from IRU-1 and IRU-2 also supplies
data to the A/T computer, the Weather Radar receiver/transmitter and the digital flight
data recorder system (DFDRS). IRU bus 3 from the left IRU provides data to the GPWS.
DAA Outputs
Heading and roll are sent to the DAAs to be converted to analog signals. Heading is
sent through the IRS transfer relays to the RDDMI for display, while roll is sent to the
yaw damper.
Digital Interface
On the following page there are three tables. One shows the digital data bus input to
the IRU. The second one shows the digital data input to the DAA. While the third shows
the digital data output from the IRU to the various systems that use IRU data and what
information is used from that data bus.
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IRS BLOCK DIAGRAM
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OPERATION
IRS TRANSFER CONTROL
Operation
The IRS transfer relays are armed with 28 volts dc from the battery bus and controlled
by a ground from the IRS transfer switch.
Discrete switching into the Weather Radar Receiver/Transmitter, EFIS Symbol Generators,
Vertical Speed Indicators, Digital Flight Data Acquisition Unit and FCCs enable the user
system to receive data from the offside IRU.
A ground through the EFIS switching relays and IRS transfer relays is used in the FCCs
to indicate when in the normal condition or if data has been transferred. An open
indicates that the FCCs will use the offside IRU for its data source. This applies to
both the local and foreign IRS transfer, and heading transfer logic.
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IRS TRANSFER CONTROL
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OPERATION
IRS DISPLAY DATA - EFIS
Operation
Each EFIS symbol generator receives digital data from each IRU. The normal data channels
are IRU 1 furnishing display data for the left EADI/EHSI and IRU 2 furnishing display data
for the right EADI/EHSI.
Either IRU may supply display data for all display data by the appropriate use of the IRS
transfer switch on the P5 overhead panel.
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IRS DISPLAY DATA EFIS
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MAINTENANCE PRACTICES
IRS BITE INTRODUCTION
Tests
The IRS built-in test equipment (BITE) provides for the ground maintenance testing of the
IRS.
IRS BITE is contained in the DAA, and is accessed through the MCDU/ANCDU/CDU. The BITE
program is divided into three fault isolation routines: current status, interface tests,
and in-flight faults storage and display.
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IRS BITE INTRODUCTION
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SYSTEM TEST
GENERAL
Interface Check
Pressing line select key 3 left (<INTERFACE CHECK), from the IRS BITE Index, will cause the
Interface Check page 1 of 4 to be displayed. The first page gives a brief description of
the test results, and allows for the operator to begin the test by pressing line select key
5 left (<TEST START). When the test start key is pressed, the Interface test begins (this
is the same test that can be performed from the IRU or the ISDU). If the Next Page key is
pressed, page 2 of 4 will be displayed. Pages 2, 3 and 4 describe the test indications that
should be displayed in the flight compartment 10 seconds after the test begins; these pages
also allow for the operator to stop the test by pressing line select key 5 left (<TEST
STOP).
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IRS BITE - INTERFACE CHECK R IRS
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IRS BITE RESULTS
GENERAL
0-2 Seconds
After the Interface Check is started, all the MSU and ISDU lights will illuminate for the
first two seconds.
2-10 Seconds
After the first two seconds, all data will fail. HDG flag on the RDDMI/RMI and on the EHSI,
OFF flag on the VSI, and ATT flag on the EADI.
After 10 Seconds
After 10 seconds the flight compartment indications should be as follows:
VSI....................................................
MAG HDG (RDDMI/RMI, EHSI) ..............................
TRUE HDG (ISDU) ........................................
PITCH ATTITUDE .........................................
ROLL ATTITUDE ..........................................
TRUE TRACK (ISDU) ......................................
MAG TRACK (EHSI)* ......................................
GROUND SPEED (ISDU) ....................................
GROUND SPEED (EADI)* ...................................
PRESENT POSITION LATITUDE (ISDU)** .....................
PRESENT POSITION LONGITUDE (ISDU)** ....................
WIND DIRECTION (ISDU) ..................................
WIND SPEED (ISDU) ......................................
WIND DIRECTION (EHSI) ..................................
WIND SPEED (EHSI) ......................................
-600 FPM
15°
10.0°
5° Up
45° Right
0.0°
22°
200 Knots
200 Knots
N22°30.0'
E22°30.0'
30° True
100 Knots
NCD
NCD
*
The data for the EFIS track and ground speed displays, when tests from the CDU are
supplied by the FMC.
**
The test results page on the CDU lists PPOS LAT as N22.5° and LONG as E22.5° which is
the same as 22°30.0'.
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IRS SELF TEST SEQUENCE
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NOTES
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IRS - SELF-TEST SEQUENCE
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SYSTEM TEST
IRU REJECTION CRITERIA
Position Error
Position error is the distance between the airplane's actual position (typically
latitude/longitude of the assigned gate) and the position that has been computed by an IRU.
The IRU position error and the time the IRU is in the NAV Mode are used to compute a point
on the position error chart. The location of this point on the chart determines if the IRU
should be removed and replaced.
If the point is above the heavily shaded area of the position error chart, the IRU should be
removed and replaced. If it is located below the heavily shaded area, the IRU is not
removed. If this point falls within the heavily shaded area of the position error chart,
the IRU should be checked again after the next flight. If after the next flight the newly
computed point still falls within the heavily shaded area of the position error chart, the
IRU should be removed and replaced.
Residual Ground Speed Error
Residual ground speed is the ground speed calculated by the IRU after a flight while the
airplane is stationary. This ground speed is displayed on the CDU on page two of POS INIT.
If this ground speed is greater than 20 knots, or greater than 15 knots on two consecutive
flights, the IRU should be removed.
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IRU POSITION ERROR CHART
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GENERAL DESCRIPTION
STANDBY HORIZON SYSTEM COMPONENT LOCATIONS
General Component Locations
Circuit Breaker
The circuit breaker supplying power to the standby horizon systems is located on the
P18-2 circuit breaker panel.
Standby Horizon Indicator
The indicator is mounted on the captain's instrument panel, P1.
STANDBY HORIZON SYSTEM COMPONENT LOCATIONS
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COMPONENT FUNCTIONAL DESCRIPTION
STANDBY HORIZON INDICATOR
1. Purpose
The standby horizon indicator is an instrument which presents pitch and roll attitude during
the entire flight of the airplane. The instrument indications are 0-360 degrees in roll and
limited to 90 degrees pitch up through 80 degrees pitch down. The instrument readings
supplement the attitude readings from other flight control cabin instruments.
2. Location
The indicator is installed on the captain's instrument panel P1.
3. Physical Description
The instrument consists of a two degree of freedom vertical gyro which mechanically controls
the horizon indicator.
The gyro spinning mass is brought to the normal rotational speed of 22,000 RPM as part of a
three phase induction motor. On the bottom of the indicator case are installed a vent to
allow air circulation and a rectangular window for viewing purposes. The indicator is
interfaced with airplane wiring through a rear mounted connector. The indicator is mounted
through three screws on the outer edges of the indicator.
4. Power
The indicator requires 28 volts dc.
5. Control
The vertical gyro is caged to airplane reference through a caging control on the front of
the indicator. Caging action is accomplished by pulling the control on the front of the
indicator. The pitch attitude presentation is adjusted five degrees pitch up or pitch down
from the caged position by rotating the caging control clockwise or counterclockwise.
6. Monitor
A flag monitors normal electrical power. Loss of power displays the flag.
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STANDBY HORIZON INDICATOR
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OPERATION
STANDBY HORIZON SYSTEM SCHEMATIC
Normal Sequence
Power
The standby horizon is powered by 28 volt dc power from the battery bus.
Monitor
The standby horizon's power-off flag is operated by a flag motor that senses presence
of power in any phase. Normally, the flag is out of view.
Operation
Following application of power, the gyro axis is caged by pulling out the alignment
knob, located on the lower right side of the instrument face.
STANDBY HORIZON SCHEMATIC
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EFIS Differences
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STANDBY HORIZON/ILS INDICATOR
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COMPONENT FUNCTIONAL DESCRIPTION
STANDBY HORIZON/ILS INDICATOR
1. Purpose
The standby Horizon/ILS indicator provides visual indications of pitch and roll attitude
information, localizer and glide slope deviation and localizer, glide slope and gyro circuit
failure with flags.
2. Features
A.
Display and Control Characteristics
Localizer needle indicates deviation from the centerline of the localizer beam.
Glide slope needle indicates deviation from the centerline of the glide slope beam.
Cage knob may be pulled momentarily after approximately 30 seconds of gyro spin up
to physically erect the gyro to the reference, vertical position.
The ILS knob provides the following:
Off - Biases the LOC and G/S needles and flags out-of-view.
ILS - Allows localizer and glide slope data to be displayed on the indicator.
B/CRS - Provides reversed signal polarity to the LOC needle for a back-course
approach to a landing field and biases the glide slope pointer from view.
The gyro stabilized ball provides a stable horizon reference.
The miniature airplane position relative to the horizon reference indicates pitch and
roll attitude.
Roll scale is displayed on the top half of the case. Pitch scale is displayed on the
face of the ball.
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STANDBY HORIZON/ILS SCHEMATIC
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OPERATION
STANDBY HORIZON/ILS SCHEMATIC
Operation
Gyro
Actuating the battery bus supplies three phase power to the standby gyro from the static
invertor. Manual erection of the gyro is accomplished by pulling the spring-loaded caging
knob.
A flag monitors normal electrical power. Loss of power displays the flag.
ILS
In the OFF mode the flags and needles are biased out of view with VHF NAV 1 tuned to an
ILS frequency and ILS mode selected, localizer and glide slope deviation drives the
needles via G1 and G2 motors.
Glide and localizer superflag voltage from the NAV receiver is used to hold the flags
out of view.
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Automatic Direction
Finder
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INTRODUCTION
AUTOMATIC DIRECTION FINDING (ADF)
1. Purpose
The ADF system is used for automatic determination of compass bearing to a transmitting
radio station, for reception of weather information and other broadcast programs, and for
flying radio ranges.
2. System Description
The ADF system is a navigational aid that receives AM radio signals and determines the
direction of the signals relative to the airplane. Signal sources utilized are standard
broadcast and low frequency radio range stations operating in the frequency range of 190 kHz
to 1750 kHz.
The ADF system is tuned and controlled from the flight compartment and receives the
electromagnetic wave transmitted by the ground station. A directional loop antenna
intercepts the magnetic portion (H) and an omnidirectional sense antenna intercepts the
electric portion (E). The ADF receiver uses these signals and airplane heading information
provided by the IRS system to compute the compass bearing of the radio station being
received. This information is converted to a synchro output that provides pointer display
on an indicator in the flight compartment. The receiver also provides audio signals to the
audio integration system.
The ADF system may also be used in a non direction finding mode in which only the sense
antenna signals are processed.
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AUTOMATIC DIRECTION FINDING (ADF)
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GENERAL DESCRIPTION
ADF SYSTEM COMPONENT LOCATION
General Component Locations
Circuit Breakers
ADF No. 1 and No. 2 circuit breakers are located on the P18 and P6 panel, respectively
(left load control center). These items are shown for location purposes only and
will not be discussed further.
Control Panel
The dual control panel is located on the aft electronics panel (P-8).
Receiver
The ADF No. 1 and No. 2 receivers are located on the electronic equipment shelf (E2-4).
Indicator
There are 2 RDDMI's installed. One is on P1 panel and the other is on the P3 panel.
Antennas
The loop antennas are located on the bottom of the airplane at station 560 and 600.
The QEC's are located above the loop antennas. The sense antennas and couplers are
located on the bottom left and right side of the airplane, just aft of the main
landing gear wheel well at station 727L, and 727R.
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ADF SYSTEM COMPONENT LOCATION
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COMPONENT FUNCTIONAL DESCRIPTION
ADF CONTROL PANEL
1. Purpose
The ADF control panel facilitates the operation of the ADF system, providing controls for
both systems.
2. Features
A.
Mode Selector Switch
The mode switches are four position switches with detents: OFF, ANT, ADF, and a
momentary spring-loaded TEST position. In the OFF position, power is removed from the
system. In the antenna (ANT) position, the receiver provides audio but not bearing
information. The ADF position provides both audio and relative bearing to the station.
The TEST position tests the receiver, providing a fixed bearing output and a test tone.
B.
Gain Control
The gain control adjusts the level of audio output signals.
C.
Tuning Controls and Frequency Display
The operator selects receiver frequency between 190 and 1749.5 in 0.5 kHz steps by using
the three concentric knobs. The outer knob controls the 1000 and 100 kHz portion of
the operating frequency. The center knob controls the 10 kHz and the inner knob
controls the 1 kHz and 0.5 kHz portion of the operating frequency. The frequency is
displayed in the frequency window.
D.
Tone Switch
The TONE switch is spring-loaded to the center OFF position and is held in the
momentary 1 or 2 position. It allows the operator to hear CW signals more clearly.
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CONTROL PANEL
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COMPONENT FUNCTIONAL DESCRIPTION
ADF RECEIVER
1. Purpose
The ADF receiver processes signals in the 190 kHz to 1749.5 kHz frequency band. The output
of the receiver represents bearing to a ground station.
2. Features
A.
Antenna Connectors
The ADF receiver has a sense antenna and a loop antenna connector located on the front
of the receiver.
B.
Test Switch
A push-button test switch provides a means of initiating a system test.
C.
Goniometer Position Indicator
A meter-like indicator shows the position of a synchro-type device within the receiver.
3. Access
The ADF receiver is secured to the shelf by a single hold down fastener. Sense and loop
antenna connectors must be disconnected before the receiver is removed from the shelf.
Additional interface with the airplane is made through an electrical connector at the rear
of the receiver.
4. Power
Both 115 volts ac and 28 volts dc are required for system operation.
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RECEIVER
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COMPONENT FUNCTIONAL DESCRIPTION
RDDMI (RADIO DIRECTION DISTANCE MAGNETIC INDICATOR)
1. Purpose
The RDDMI displays ADF or VOR bearing information. The display function is selected by
operating ADF/VOR mode selector switches.
2. Features
A.
Bearing Pointers
The No. 1 radio bearing pointer is the single line pointer that displays VOR-1 or ADF-1
bearing information.
B.
Pointer Mode Switches and Annunciators
The pointer mode switches are push-on, push-off types and determine whether the
pointers are controlled by ADF or VOR information. When the pointer is switched to the
ADF mode, the ADF legend and the associated arrow come into view. The arrows are
miniature representations of their respective radio pointers. The legend color is the
same as the color of the radio pointers.
C.
Pointer Warning Flags
The flags come into view in the absence of valid input data from the selected ADF or
VHF NAV system.
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RDDMI
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COMPONENT FUNCTIONAL DESCRIPTION
LOOP ANTENNA AND QUADRANTAL ERROR CORRECTOR
1. Purpose
A.
Loop Antenna
The loop antenna is a direction sensitive RF pickup device designed to accept the
magnetic portion of the received electromagnetic wave for use in the ADF system.
B.
Quadrantal Error Corrector
The quadrantal error corrector compensates for the distortion introduced to the loop
antenna's reflection pattern by the airplane's wings and fuselage. The device then
couples the RF signals from the loop antenna to the antenna cable.
2. Location
Quadrantal Error Corrector
The corrector is mounted directly on top of the loop antenna assembly.
3. Features
Loop Antenna
The loop antenna is a sealed, non-rotating antenna consisting of two pairs of ferrite
core coils. One pair is aligned parallel to the longitudinal axis of the airplane; the
other pair aligned parallel to the lateral axis. The coils interface with the ADF
system through a single connector on the top of the antenna.
4. Access
A.
Loop Antenna
Access to the loop antenna is gained by first removing eight screws which fasten the
antenna cover to the airplane.
B.
Quadrantal Error Connector
The corrector is accessed by first removing the loop antenna.
5. Maintenance Practices
Loop Antenna
Before installing the antenna, assure that the mating surfaces of the antenna and
airplane structure are clean to provide a low resistance electrical bond. A coating of
corrosion preventive compound should be applied to the mating surfaces as outlined in
the Maintenance Manual.
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LOOP ANTENNA AND QUADRANTAL ERROR CORRECTOR
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COMPONENT FUNCTIONAL DESCRIPTION
SENSE ANTENNA AND COUPLER
1. Purpose
A.
Sense Antenna
The sense antenna is an omnidirectional RF pickup device designed to accept the electric
portion of the received electromagnetic wave for use in the ADF system. The sense
antenna provides high strength signals to the ADF receiver.
B.
Coupler
The sense antenna coupler matches the capacitance of the ADF sense antenna to the
capacitance of the transmission line between the antenna and the receiver.
2. Location
A.
Sense Antenna
The sense antenna is installed on the left wing to body fairing, aft of the aft wheel
well bulkhead.
B.
Coupler
The coupler is mounted to the aft bulkhead of the left main wheel well.
3. Features
A.
Sense Antenna
The sense antenna is a portion of the reinforced wing to body fairing that has been
flame sprayed with aluminum. This metal coating is used as the antenna element. The
area is painted with an antistatic paint which provides precipation static protection.
The signal is coupled from the antenna by a feed stud.
B.
Coupler
The coupler contains the necessary circuitry to provide optimum ADF system performance
and has a single connector to attach the triaxial cable to the ADF receiver.
4. Maintenance Practices
Sense Antenna
The metal coating that is used as the antenna is not removable from the fairing. Sense
antenna replacement is accomplished by replacing the wing to body fairing as detailed in
Chapter 53 of the Maintenance Manual.
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SENSE ANTENNA AND COUPLER
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OPERATION
ADF SYSTEM - BLOCK DIAGRAM
Operation Sequence
Power
ADF System 1 receives 28 volts dc from the standby bus and ADF system 2 receives 28
volts dc from the elex-2 bus. Synchro excitation for the systems is 26 volts ac
from instrument transformers number 1 and 2.
Signal Processing
The loop antenna connects the magnetic portion of the RF signal to the receiver through
the quadrantal error corrector (QEC).
The sense antenna connects the electric portion of the RF signal to the receiver
through the sense as antenna coupler. The coupler matches the 3000 PF capacitance of
the receiver to the antenna.
From the control panel, the operator selects the receiver frequency, the desired mode
of operation enables the tone oscillator for CW operation and tests the receiver.
Modes
During the ADF mode of operation, both antenna inputs are used by the receiver to
develop a bearing signal. ADF receiver 1 controls the narrow pointers on each RDDMI.
The pointer automatically points in the direction of the selected ground station.
Audio, which is received through the sense antenna, is also available to the audio
selector panels through the gain control on the ADF control panel.
During the ANT mode, only the sense antenna is used and provides reception of audio to
the audio selector panels. The RDDMI pointers are driven to the 9 o'clock position.
Test
A test of the system is initiated at the control panel or at the receiver.
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ADF SYSTEM BLOCK DIAGRAM
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NOTES
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EFIS Differences
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COMPONENT FUNCTIONAL DESCRIPTION
ADF DISPLAYS - EHSI
Purpose
The EHSI may be used to display ADF bearing in all modes.
Features
ADF Bearing Pointers
ADF bearing pointers and reciprocals, indicating direction to and from selected ADF
stations, are shown in all modes.
ADF Bearing Vectors
ADF bearing vectors may be displayed in MAP or CTR MAP modes by pressing the VOR/ADF
switch on the EFIS control panel. ADF bearing vectors point from the airplane symbol
to the selected ADF stations.
ADF Fault Displays
The ADF pointer or vector will be removed if an invalid condition is indicated by a
loss of ADF power.
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ADF DISPLAYS - EHSI (TYPICAL)
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OPERATION
ADF SYSTEM - BLOCK DIAGRAMA
Operation Sequence
Power
ADF System 1 receives 28 volts dc from the STBY bus and ADF system 2 receives 28 volts
dc from the Elex-2 bus via the Dual ADF Control Panel. Synchro excitation for the
systems is 26 volts ac from ADF-1 and ADF-2 Excitation Transformers.
Signal Processing
The loop antenna connects the magnetic component of the rf signal to the receiver
through the quadrantal error corrector (QEC). The sense antenna connects the electric
component of the RF signal to the receiver through the sense antenna coupler. The
coupler matches the 3000 PF capacitance of the receiver to the antenna.
The control panel provides mode and frequency selection as well as selection of the
1020 Hz tone.
Modes
During the ADF mode of operation, sense and loop antenna inputs are used by each
receiver to develop a bearing signal. The computed bearing is sent to the RDDMI's and
to the EFIS symbol generators for display on the EHSI's. ADF receiver 1 controls the
narrow pointer and ADF receiver 2 controls the WIDE pointer on each RDDMI and EHSI.
The pointer automatically points in the direction of the selected station. Audio from
the receiver is available to the digital audio control system through the gain control
on the ADF control panel.
During ANT mode, only the sense antenna is used.
EHSI are driven to the 9 o'clock position if ANT
bearing vectors will likewise be driven to the 9
modes if the VOR/ADF switch is pressed. In all
ADF valid for the RDDMI and EHSI.
The ADF pointers on the RDDMI and the
mode is selected. In ANT mode the
o'clock position in the MAP/CTR MAP
modes, the ADF ON signal is used as the
Test
A test of the system may be initiated at the control panel or at the receiver. When
test is selected, the pointers on the EHSI and the RDDMI will point to 315° relative to
the centerline of the airplane and the goniometer position indicator on the receiver
will point to the 45° test position.
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ADF System Block Diagram
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MAINTENANCE PRACTICES
ADF SYSTEM - MAINTENANCE SUMMARY
1. Fault Isolation
Refer to the Maintenance Manual, Chapter 34, Section 57, for all troubleshooting and testing
procedures.
2. Remove/Install
A.
General
Refer to the Maintenance Manual, Chapter 34, Section 57 for component removal and
replacement details.
B.
Loop Antenna
Ensure mating surfaces between the antenna and airplane meet electrical bonding
requirements.
3. Test Preparations
A.
Precautions
Make certain that the area around the ADF antenna is clear of metal objects such as
baggage carts, fuel trucks, fuel pits, tools, etc. Metal objects on the ramp can
distort the radio signal and result in ADF bearing error.
B.
Power
The 115 volt ac and 28 volt dc buses must be energized. Circuit breakers for the ADF,
IRS and interphone systems must be closed.
4. Functional Test
A.
Control Panel
Pressing and holding the test switch causes the ADF pointer to indicate to 315°
relative bearing and a clear audio tone.
Selecting ADF mode causes the ADF pointer to indicate correct magnetic heading to
the selected station.
B.
Receiver
Pressing the test switch on the receiver causes the goniometer position indicator to
rotate to 45 degrees. Releasing the test switch causes the indicator to return to its
normal position for the mode selected on the control panel.
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VHF Navigation
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
VHF NAVIGATION SYSTEM
1. Purpose
The VHF navigation system provides airplane position data for in-route navigation and for
landing approach.
2. System Description
The VHF navigation system performs 2 types of operation, VOR and ILS.
VHF omnidirectional range (VOR) is designed to provide accurate in-route navigation. In VOR
operation, the system supplies bearing and deviation with respect to a VOR station.
The instrument landing system (ILS) provides vertical and horizontal guidance when making an
approach to a runway. In ILS operation, the system supplies deviation from the localizer
(LOC) and glide slope (G/S) beams.
The basic VHF navigation system is composed of receivers and processors an antenna system
comprised of a glide slope antenna, a localizer antenna and a VOR antenna and a control
panel.
The antennas supply RF inputs which the receivers processes to supply outputs to interfacing
components. The VOR antenna supplies RF during cruise operation. The localizer and glide
slope antennas supply RF during approach. The receiver is manually tuned from the control
panel and automatically from a flight management system. The omnibearing output of the
receiver is applied to the flight management system for present position correction.
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VHF NAVIGATION SYSTEM
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
VHF NAVIGATION SYSTEM COMPONENT LOCATION
General Component Locations
Circuit Breakers
VHF NAV-1 circuit breakers are located on the P18-2 panel (left load control center).
The circuit breakers for VHF NAV-2 are located on the P6-1 panel (right load control
center).
Control Panels
There is one control panel for each system. Both are located on the aft electronics
panel (P8). The No. 1 control panel is on the left and the No. 2 panel is on the
right.
Course select controls are located on DFCS mode control panel on the P7.
NAV Units
The VHF NAV receiver units No. 1 and No. 2 are located on the E3-4 shelf.
Antennas
The VOR/LOC antenna is located on top of the vertical stabilizer. The localizer and
glideslope antennas are on the forward bulkhead under the nose radome.
Transfer and Interface Components
There is a VOR/ILS - NAV transfer switch on the left and right side of panel P7. The
VHF NAV transfer switch is located on the overhead panel P5.
VHF NAV unit transfer relays and ILS antenna transfer relays are located on the E3-4
shelf. The flight instrument accessory unit and digital/analog adapter are located on
the E1-2 shelf.
Indicators
An ADI, HSI, and RDDMI are located on each pilot's panel, P1 and P3
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MAINTENANCE TRAINING MANUAL
VHF NAVIGATION SYSTEM COMPONENT LOCATION
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COMPONENT FUNCTIONAL DESCRIPTION
VHF NAV/DME CONTROL PANEL
1. Purpose
The panel provides controls for the selection of automatic or manual frequency control for
the VHF NAV and DME systems. Segmented displays are provided for both the manual and
automatic frequency selections. Switches are provided to initiate self-test of the VHF NAV
and DME systems.
2. Physical Description
The AUTO/MAN mode selector switch is a push-on/push-off illuminated switch that is used to
select either AUTO or MAN frequency control of the VHF NAV and DME systems. The selected
frequency control mode is annunciated by the switch. The annunciator legends are white on
black background, and are visable only when the internal lamps are illuminated. AUTO
selection is armed by the airplane NAV/VOR-ILS switching. NAV must be selected for display
on the HSI before AUTO frequency control can be enabled.
Dual concentric controls are
knob controls tens and units
frequency range is 108.00 to
selected is displayed by the
provided for the manual selection of the frequency. The outer
and the inner knob controls tenths and hundredths. The
117.95 MHz, selectable in 50 kHz steps. The frequency manually
manual frequency display.
The MANUAL frequency display shows the manual frequency selected. When the system is in the
AUTO mode, a dayglo fire orange FLAG (bar) is raised over the manual display window, and the
frequency selected by the flight management system appears in the AUTO display. The
frequency range is extended to 135.95 MHz during AUTO mode.
With the system in the manual mode, the flag is retracted from the manual display and the
automatic display is blanked.
The frequency displays and mode annunciator are tested through the airplane master test
circuit. When the master test switch is actuated, both displays read 188.88, flashing 2
seconds on and 1 second off. Both the AUTO and MAN annunciations are illuminated steadily.
3. Self-Test
Two switches, three position, spring loaded to the center position, are provided to selftest the VHF NAV system. The VOR test initiates a test of the VOR receiver within the NAV
unit. The ILS test initiates a test of the localizer and glideslope receivers within the
NAV unit. Activation of a VOR, UP/LT, or DN/RT test interrupts the flight management system
interlock circuit.
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MAINTENANCE TRAINING MANUAL
VHF NAV/DME CONTROL PANEL
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
VHF NAVIGATION UNIT
1. Purpose
The unit is an airborne receiver designed to process ground station transmission signals in
the very high and ultra high frequency range. The outputs of the unit are TO-FROM, relative
bearing, VOR, localizer and glide slope signals. In addition, the unit supplies warning
outputs for monitor purposes of receiver operation.
2. Physical Description
The unit is housed in a standard ARINC 404A, 1/2 ATR short case, with hinged panels. It is
interfaced with airplane wiring through a rear-mounted, two-section connector. Another
connector, in the upper rear, provides test connections. The unit is mounted in the
airplane through two sets of hold-down hooks and lock units. A handle is provided for
handling purposes. The receiver operates on two bands of frequencies, 108.00 MHz to 117.95
(VHF) with 50 - KHz spacing and 329.15 MHz to 335 MHz (UHF) with 150 KHz spacing. The VHF
band provides 160 channels of VOR and 40 channels of LOC operation. The VHF operation
provides 40 channels for glide slope, each channel coupled to a corresponding LOC channel.
VHF NAVIGATION UNIT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
VOR BEARING DISPLAY
1. Purpose
VOR bearing is displayed on the radio direction distance magnetic indicator (RDDMI). The
RDDMI also provides the pilot with compass heading, ADF bearing, and DME distance. Only the
VOR bearing display is discussed in the following text.
2. Physical Description
The RDDMI displays ADF or VOR bearing information. The display depends upon whether VOR or
ADF data is selected through the ADF/VOR selector switches. The No. 1 radio bearing pointer
displays VOR-1 or ADF-1 bearing information and the No. 2 pointer displays VOR-2 or ADF-2
data. The ADF/VOR switches are push-on, push-off types and control arrow symbols that
annunciate either ADF or VOR legends. The legends ADF/VOR and the associated arrow are
lighted by the instrument lighting system. The arrows are miniature representations of
their respective radio pointers and the legend color of the arrows is the same as the color
of the radio pointers. The bearing pointers are driven by synchro receivers which are
controlled by synchro transmitters in the ADF or VOR receiver. The radio flags come into
view in the absence of valid input data from the selected VHF NAV or ADF system.
VOR BEARING DISPLAY
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
HORIZONTAL SITUATION INDICATOR
1. Purpose
The HIS displays navigation information from either the VHF NAV unit or the FMC. Displays
associated with the VHF NAV system are discussed here.
2. Physical Description
A.
NAV - VOR/ILS Annunciator
VOR/ILS is annunciated when the instrument is displaying VHF NAV information. The
legend NAV appears whe FMC information is displayed.
B.
NAV Flag Annunciator
An orange and yellow striped warning flag appears in the course mask when the NAV
display is not valid.
C.
Course Pointer
Selected course and its reciprocal is displayed by a dagger-shaped cursor, located in
the center of the heading card when VOR/ILS is displayed. Selected course is read
against the heading card while course error is read with respect to the lubber line.
The dagger and pointer serve as the index for lateral deviation indication.
D.
TO/FROM Indicator
An indicator points toward or away from the course select or to show that the airplane
is flying to or away from a VOR station.
E.
Course Deviation Scale and Deviation Bar
Deviation from a selected course is indicated by a bar across a scale affixed to the
rotating mask. Scale markings are four white dots, two on either side of the scale
center. Scale values are one degree per dot for LOC and 5 degrees per dot for VOR.
F.
Vertical Deviation Scale, Pointer, and Flag
Deviation from the glideslope path is indicated b means of a pointer which moves beside
a vertical scale. The pointer and the VERT flag are held out of view in the VOR mode.
In the ILS mode, the GS deviation scale is 0.35 degrees per dot. An orange VERT flag
obscures the scale for invalid displays.
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MAINTENANCE TRAINING MANUAL
HORIZONTAL SITUATION INDICATOR
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COMPONENT FUNCTIONAL DESCRIPTION
ATTITUDE DIRECTOR INDICATOR
1. Purpose
The attitude director indicator (ADI) displays information from the flight management system
and the VHF NAV unit. Only the displays driven by the VHF NAV unit are addressed here.
2. Physical Description
Glide slope deviation is displayed by reading the pointer against the glide slope scale.
The GS flag is in view during a malfunction of the glide slope portion of the VHF NAV unit.
Localizer deviation is displayed by reading the runway symbol against the scale at the
bottom of the indicator.
The RUNWAY flag appears if the runway or its drive system malfunction.
ATTITUDE DIRECTOR INDICATOR
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
VOR ANTENNA
1. Purpose
The VOR antenna is designed to receive very high frequency radio signals between the
frequencies of 108.00 and 118.00 megahertz.
2. Physical Description
The antenna
composed of
the leading
A lightning
is a dual element antenna with a
two balanced loops enclosed in a
and training edges and the whole
diverter strip is embedded along
characteristic impedance of 50 ohms. It is
fiberglass housing with metallic fin tips on
assembly is located on top of the vertical fin.
the top edge of the housing for protection.
3. Maintenance Practices
To remove the antenna, first remove the 12 leading edge fin tip mounting screws. Then
remove the leading edge fin tip. Remove the two rf connectors from antenna. Remove the 40
antenna mounting screws along the base of the antenna. Do not remove the eight screws
securing the trailing edge fin tip to antenna. Remove the antenna and trailing edge fin
tip. Reassemble in reverse order following the procedure in the Maintenance Manual.
VOR/LOC ANTENNA
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
LOCALIZER ANTENNA
1. Purpose
The localizer antenna is designed to receive very high frequency radio signals between 108
and 112 megahertz.
2. Physical Description
The antenna is horizontally polarized and has a characteristic impedance of 50 ohms. It is
a dual-element antenna, terminating in a power divider and two female connectors mounted on
the left through two relays. The antenna is mounted below the weather radar antenna through
six screws on each side.
LOCALIZER ANTENNA
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
GLIDE SLOPE ANTENNA
1. Purpose
The glide slope antenna is designed to receive ultra high frequency radio signals between
329.0 and 335 megahertz.
2. Physical Description
The antenna is a single, horizontally-polarized unit and has a characteristic impedance of 50
ohms. There are two coaxial connectors for routing the incoming rf signals to the VHF
navigation units. The antenna is mounted to the bracket with six screws.
GLIDE SLOPE ANTENNA
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
INSTRUMENT SWITCHING PANELS
1. Purpose
The HSI-VOR/ILS transfer switch selects the data source for display on the HSI. The threeposition VHF NAV transfer switch is used to select the VHF NAV unit controlling sources the
pilots' instruments and other airplane systems.
2. Physical Description
Separate switches are provided for the captain's and the first officer's HSIs. Only the
captain's switch is shown. With the switch in the VOR/ILS position, lateral and vertical
deviation and TO/FROM inputs are provided from the VHF navigation system. In the NAV
position, the input data is provided by the flight management system.
In the NORMAL position, the captain's instruments are driven from the VHF NAV unit-1, and the
first officer's instruments are driven from the VHF NAV unit-2. When a switch is placed in
the BOTH ON 1 or BOTH ON 2 position, both pilots' instruments are controlled form the
selected VHF NAV unit.
INSTRUMENT SWITCHING PANELS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
COURSE SELECT CONTROLS
1. General
The course select controls and indicators are part of the DFCS mode control panel, located
on the glare shield (P7).
2. Physical Description
The number one control provides VOR or LOC course selection for the VHF NAV unit-1, preset
course for the flight control computer-A, and positions the course cursor on the captain's
HSI. The course selected is displayed in the solid state liquid crystal indicator located
above the course select control. The number two control provides the same inputs for VHF NAV
unit-2, FCC-B, and the first officer's HSI.
COURSE SELECT CONTROLS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
FLIGHT INSTRUMENT ACCESSORY UNIT
1. Purpose
The flight instrument accessory unit contains components for interfacing the VHF navigation
system with other airplane navigation systems.
2. Physical Description
The flight instrument accessory unit is a Boeing manufactured component. The components
contained in the unit include resistors, diodes, relays, dimming circuits, and step down
transformers. The following is a list of the relays and the function performed by each:
K17
K16
Agility RLY-1
Agility RLY-2
K12
K14
VOR/ILS Test
Inhibit
K13
K15
Ant. Transfer
Relays
The unit is installed through the action of a camlock handle and is interfaced with airplane
wiring through a dual connector mounted in the rear of the unit.
FLIGHT INSTRUMENT ACCESSORY UNIT
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MAINTENANCE TRAINING MANUAL
NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
VOR OPERATION AND THEORY
Operation
SHEET 1
VOR ground stations transmit on frequencies in the range of 108.00 and 117.95 MHz,
except on odd tenths up to 111.90 MHz (which is reserved for localizer channels).
The principle of VOR navigation involves a measurement of phase difference between two
transmitted signals, one termed the reference signal, the other the variable signal.
The reference signal is an RF carrier amplitude modulated with a 9960 Hz subcarrier.
The subcarrier is in turn frequency modulated ±480 Hz at a 30 Hz rate. The signal is
radiated by an omnidirectional antenna, and the phase of the received signal is the
same on all VOR radials.
The variable signal is an unmodulated RF carrier (CW) directional signal which is
rotated through 360 degrees at 30 rps. The resulting received signal fluctuation
creates the same effect in the receiver's detector as would a 30 Hz amplitude modulated
transmitted carrier.
The reference 30 Hz signal and the variable 30 Hz signal are in phase when the variable
rotates past (is aligned with) the magnetic north (000) radial. An airplane located on
the radial receives both signals in phase. On any other radial, the phase difference
is equal to the angle of the radial from magnetic north. Processing of the phasedifference signals results in output signals that are used to position VOR bearing
pointers in the RMIs and HSIs.
SHEET 2
The principle of VOR deviation involves determining the difference between selected
course and VOR bearing. VOR magnetic bearing sine and cosine signals are combined with
selected course information. The difference between these two parameters is the VOR
deviation. For a difference of 5 degrees right or 5 degrees left, the receiver output
is standardized at 75 microamperes which causes a VOR deviation bar deflection of one
dot. For a difference of 10 degrees, the receiver output is 150 microamperes which
produces a two-dot deflection.
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MAINTENANCE TRAINING MANUAL
VOR OPERATION AND THEORY (SHEET 1)
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MAINTENANCE TRAINING MANUAL
VOR DEVIATION PROCESSING
VOR OPERATION AND THEORY (SHEET 2)
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
LOCALIZER THEORY
Operation Sequence
The localizer signal is transmitted on frequencies in the range of 108.1 to 111.9 MHz (odd
tenths).
The principle of localizer operation involves a comparison of 2 transmitted signals: one
modulated at 90 Hz, and the other modulated at 150 Hz.
These signals are radiated to produce two directional lobes, one on either side of the
runway centerline. The left lobe is modulated with the 90-Hz signal and the right lobe with
150 Hz. The on-course path is a line where the two audio signals are equal. It coincides
with the centerline of the runway.
When the airplane position is to the left of the on-course path, the 90-Hz signal
predominates and localizer deviation indicators are deflected to the right, indicating that
the runway centerline is to the right. To the right of the on-course path, the 150 Hz
signal predominates and the localizer deviation pointers deflect to the left. On a flight
path one degree left or right of the on-course path, the receiver output is standardized at
75 microamperes which causes a localizer deviation bar deflection of one dot. On a line
displaced two degrees, the receiver output is 150 microamperes which produces a two-dot
deflection.
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LOCALIZER THEORY
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MAINTENANCE TRAINING MANUAL
OPERATION
GLIDE SLOPE THEORY
Operation Sequence
The glide slope signal is transmitted on UHF channels on frequencies in the range of 329.3
to 335.0 MHz. Glide slope receiving circuits are tuned automatically whenever a localizer
frequency is selected.
The principle of glide slope operation involves a comparison of 2 transmitted signals: one
modulated at 90 Hz, and the other modulated at 150 Hz.
These signals are radiated to produce two lobes, one above the other. The upper lobe is
modulated with 90 Hz, the lower lobe with 150 Hz audio. The glide slope is laterally
aligned with the localizer path on a line 2.5 to 3 degrees above ground level where the two
audio signals are equal.
When the airplane position is above the glide slope, the 90 Hz signal predominates and the
glide slope pointers are deflected down indicating that the glide slope is below the
airplane. Below the glide slope, the 150 Hz signal is strongest and the glide slope
pointers deflect upwards indicating that the glide slope is above the airplane. On flight
paths 0.35 degrees above or below the glide slope, the receiver output is standardized at 75
microamperes which produces one dot of glide slope needle deflection. On a path displaced
0.7 degrees, receiver output is 150 microamperes which results in a two-dot deflection.
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MAINTENANCE TRAINING MANUAL
GLIDE SLOPE THEORY
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MAINTENANCE TRAINING MANUAL
OPERATION
VHF NAVIGATION SYSTEM BLOCK DIAGRAM
Operational Sequence
The block diagram shows the interface between the VHF NAV system components and between the
VHF NAV system components and other airplane systems.
Two 28 volt dc and 26 volt ac sources are required to operate each VHF NAV unit. The 26
volt ac is also used as reference for the RDDMI VOR bearing pointers and for the MAG HDG
synchros (TX).
The VHF NAV units are tuned manually from the control panels or are tuned automatically
through the DAAs when NAV mode of operation is selected for the HSIs. The units may also be
tested from the control panels through the flight instrument accessory unit, provided the
digital flight control system (DFCS) is not armed for VOR or ILS operation.
During VOR operation, the VOR/LOC antenna supplies RF inputs to the VHF NAV units. When the
DFCS is armed for ILS operation, the LOC antenna replaces the VOR/LOC antenna and the G/S
antenna supplies VHF signals to the VHF NAV units.
The RDDMIs are the sources of magnetic heading and the DFCS is the course reference sources
for the VHF NAV units.
When VOR frequencies are tuned for the VHF NAV units, VOR/LOC deviation and superflag are
supplied to the HSIs during VOR/ILS operation, the DAAs and the FCCs, TO-FROM is supplied to
the HSIs, VOR bearing is supplied to the RDDMIs, except during DME agility turning, and SINCOS VOR bearing is supplied to the DAAs for position update. Audio from the VHF NAV units is
supplied into the audio integrating system. During VOR operation, the expanded LOC
indicator in ADI and the G/S pointers in the ADI and HIS or stowed out of view.
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MAINTENANCE TRAINING MANUAL
OPERATION
Localizer Operation
When LOC frequencies are tuned for the VHF NAV units, LOC signals are received by the
VOR/LOC antenna or, when the approach mode is selected for the DFCS, by the LOC antenna
located in the radome. The localizer signal is processed by the NAV unit and LOC
deviation, which is the deviation from the localizer course center line, is produced.
LOC deviation are applied to the instrument comparator, FCCs, and to the DAAs and is
provided to the HSI and ADI. The VOR/LOC superflag signal is provided to all equipment
using the LOC deviation signal.
When localizer frequencies are selected, ILS DC logics are provided to the comparator,
FCCs, the DAAs and the ADI, enabling processing and display of the LOC deviation
signal.
Glide Slope Operation
The glide slope operation is enabled when localizer frequencies are selected. The
glide slope signal is received by a glide slope antenna, located in the radome. The
signal is processed by the VHF NAV units, which compute the airplane displacement from
the center line of the glide slope path, and is a measure of GS deviation.
The glide slope deviation signal are provided by the VHF NAV units directly to the
instrument comparator and FCCs Deviation is also provided to the ADI, GPWS computer and
the HSI.
The GS valid signal is also provided to all equipment receiving the deviation signal.
Switching
During abnormal operation, a pilot may switch his instruments and associated components
from the normal VHF NAV system to the alternate system by operating the VHF NAV
transfer switch. The VOR/LOC superflag signals are applied through a series of switch
and relay contacts to monitor proper switch and relay operation.
VHF Navigation Self-Test
Navigation system self-test is enabled through switches on the control panel. Switches
are provided to initiate a VOR, ILS UP/L, and ILS DN/R test.
A VOR frequency must be selected to arm the VOR test and a LOC frequency must be
selected to arm the ILS tests. When the approach mode is selected for the DFCS, all
navigation system self-tests are inhibited.
The ILS UP/L self-test is also initiated during the VHF NAV interface test of the DFCS
BITE operation.
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MAINTENANCE TRAINING MANUAL
OPERATION
VHF NAV SYSTEM SWITCHING SCHEMATICS - VOR/ILS
Operation
VOR/ILS Switching
VHF NAV Transfer Switching
Placing the VHF NAV transfer switch to the BOTH ON 1/BOTH ON 2 VOR positions
supplies 28 volts dc through contacts of VHF NAV transfer switch and through
closed contacts of transfer relays 1 or 2 driving the relay to the requested
position. Returning the switch to NORMAL drives the appropriate relay to the
original position.
HSI Transfer Switching
The HSI transfer switches are two position switches. Placing an HSI transfer
switch in NAV position, steps the associated VOR/ILS - NAV transfer relay to the
NAV position.
The HSI transfer circuits are transferred to the VOR/ILS position selelecting the
VOR/ILS position on the HSI transfer switch.
Antenna Switching
During VOR operation the fin tip VOR/LOC antenna supplies rf through the ILS relays and
the VOR rf transfer relays to the navigation units. The nose radome localizer antenna
is selected through the ILS relays by mode logic in the DFCS flight computers when
making an ILS approach. The logic requires that an ILS frequency be tuned; DFCS
approach mode must be selected; and the VOR/ILS test inhibit relays in the flight
instrument accessory unit energized. ILS logic is also supplied to the yaw damper
computers for gain programming.
The glide slope antenna elements are connected through the glide slope rf transfer
relays to the navigation units.
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VHF NAV SWITCHING SCHEMATIC - VOR/ILS
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NOTES
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VHF NAV SWITCHING SCHEMATIC - ANTENNAS
MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
VOR RECEIVER SELF-TEST
Functional Test
The VOR test checks the operation of the VOR receiver within the NAV unit. Test results are
displayed on the appropriate indicators.
The VOR test preparation and procedure described here is for the VOR/ILS navigation system
No. 1. The test of system No. 2 is performed in the same way, using the controls and
indicators for system No. 2.
Test Preparation
Prior to testing system No. 1, the following switches must be in the position
indicated:
VHF NAV Transfer Switch................ NORMAL
Capt's HSI Transfer Switch............. VOR/ILS
Course Select No. 1.................... 000
Next, select a VOR frequency of 108.00 MHz on the VHF NAV/DME-1 control panel. Observe
the VOR status on the sensor status display on the flight management computer control
display unit (FME CDU).
Test Procedure
To initiate VOR test, press and hold the test switch in the VOR position, and observe
the following displays:
Captains's HSI
Deviation Bar ......................... CENTERED
To/From Flag .......................... FROM
Failure Flags ......................... SEE NOTE
RDDMIs
No. 1 Pointer (VOR).................... 180° (±3°)
Failure Flags ......................... SEE NOTE
FMC CDU
VHF NAV Status ........................ OK
NOTE:
Warning flags remain in view for 3 seconds; retract from view for 20 seconds;
reappear 23 seconds after initiating test for as long as self-test is
actuated. Times are approximate.
Release the test switch and all VOR indications return to normal.
Repeat the VOR test with the VHF NAV transfer switch in the BOTH ON 1 position and BOTH
ON 2 position. Test displays appear on the captain's and first officer's instruments
and the instrument comparator LOC and G/S lights are inhibited.
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VOR RECEIVER SELF TEST
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
ILS RECEIVER SELF-TEST
Functional Test
The ILS test preparation and procedure described here is for the No. 1 system. The best of
the No. 2 system is performed in the same way, using the No. 2 controls and indicators.
Test Preparations
Prior to testing system No. 1, the following switches must be in the indicated
positions.
VHF NAV Transfer Switch................ NORMAL
CAPT's HSI Transfer Switch............. VOR/ILS
Next, select a localizer frequency of 108.10 MHz on the VHF NAV/DME-1 control
panel. Observe the ILS status on the Sensor Status display on the FMC CDU.
Test Procedure
To initiate the ILS test, press and hold the test switch in the UP/LT position, and
observe the following displays:
Captain's HSI
Deviation Bar ......................... LEFT 1 DOT
G/S Deviation ......................... UP 1 DOT
Failure Flags ......................... OUT OF VIEW
Captain's ADI
Expanded Localizer..................... LEFT 1 DOT
G/S Deviation ......................... UP 1 DOT
Failure Flags ......................... OUT OF VIEW
RDDMI's
Pointer 1 ............................. 3:00 O'clock Position
FMC CDU
VHF NAV Status ........................ OK
Release the test switch and all ILS indicators return to normal.
Perform the test again with the switch in the DN/RT position. The deviation bar and
expanded localizer drive right 1 dot, and the glide slope deviation pointer drives down
1 dot.
Repeat the ILS test with the VHF NAV transfer switch in the BOTH ON 1 position and BOTH
ON 2 position. Test displays appear on the captain's and first officer's instruments,
and the instrument comparator LOC and G/S lights are inhibited.
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MAINTENANCE TRAINING MANUAL
ILS RECEIVING SELF TEST
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NOTES
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EFIS Differences
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COMPONENT FUNCTIONAL DESCRIPTION
VOR/ILS NAV SYSTEM CONTROL
Purpose
The EFIS control panel provides mode selection to the EHSI.
The three-position VHF NAV transfer switch selects which VHF NAV unit will be used by the
EFIS symbol generator to drive the EHSI displays.
Physical Description
The EFIS control panel mode select switch provides selection of the VOR/ILS mode on the
EHSIs. The VOR/ILS mode can be displayed in the FULL compass rose format or the EXPANDED
compass rose format.
In the MAP or CTR MAP modes, VOR vectors can be displayed by pressing the VOR/ADF map
selection switch on the EFIS control panel.
With the VHF NAV transfer switch in the NORMAL position, VHF NAV unit-1 provides information
to the captain's instruments, and VHF NAV unit-2 provides information to the first officer's
instruments. When the transfer switch is placed in the BOTH ON 1 or BOTH ON 2 positions,
both the captain's and first officer's instruments are controlled by the selected VHF NAV
unit.
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MAINTENANCE TRAINING MANUAL
VOR/ILS NAV SYSTEM SWITCHING
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
ELECTRONIC HORIZONTAL SITUATION INDICATOR
Purpose
The EHSI displays navigational information from the VOR/ILS NAV system or from the flight
management system.
Physical Description
Mode Annunciator
VOR or ILS along with source (1 or 2) are annunciated.
Deviation Bar
Deviation from a selected course is indicated by the displacement of a bar across a
scale centered on the airplane symbol. Scale markings are four dots. The scale
represents one degree per dot for LOC and 5 degrees per dot for VOR.
Course Pointer and TO/FROM
Selected course is displayed by a dagger-shaped pointer in the full VOR/ILS mode and by
a small pointer with a line extending to the HDG scale in the exp VOR/ILS mode.
TO/FROM is annunciated by a pointer between the airplane symbol and the selected course
pointer in the full mode, and by the word TO or FROM in the lower right corner of the
display in both the full and expanded modes.
G/S Deviation Display
G/S deviation is displayed in ILS mode on a scale at the right of the EHSI display.
The index is a truncated triangle and the scale is 0.35 degrees per dot.
Frequency Display
VOR or ILS frequency as selected on the VHF NAV control panel is displayed in the lower
right corner. If AUTO is selected on the VHF NAV control panel, the EHSI frequency
display will read AUTO.
VOR Vectors
The VOR vectors are available when the EFIS control panel mode selector is in CTR MAP
or MAP position and the VOR/ADF MAP selection switch is on.
The VOR vectors will be
blank when the FMC initiates agility tuning.
Data Anomalies
Displays for no-computed data and invalid data for each input parameter are shown in
the table.
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MAINTENANCE TRAINING MANUAL
ELECTRONIC HORIZONTAL SITUATION INDICATOR - VOR DISPLAYS
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NOTES
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MAINTENANCE TRAINING MANUAL
ELECTRONIC HORIZONTAL SITUATION INDICATOR - ILS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
ELECTRONIC ATTITUDE DIRECTOR INDICATOR
Purpose
The electronic attitude director indicator displays localizer and glide slope deviation from
the VHF NAV receivers.
Physical Description
Glide slope deviation is displayed on the right side of the EADI, and localizer deviation is
displayed on the bottom of the EADI.
On the ±2-dot standard localizer deviation scale, each dot of deviation represents
approximately one degree. When the deviation is less than 5/8 of a dot and the VOR/LOC mode
is selected on the DFCS mode control panel, the scale automatically changes to the expanded
localizer scale in which one dot represents approximately one-half degree.
Indications and annunciations for no computed data and/or invalid data are shown.
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MAINTENANCE TRAINING MANUAL
ELECTRONIC ATTITUDE DIRECTOR INDICATOR - ILS DISPLAYS
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
VHF NAVIGATION SYSTEM-1 BLOCK DIAGRAM
Operation Sequence
The block diagram is shown for system one; system two is identical.
Power for the VHF NAV system is supplied by the P18 circuit breaker panel. 26 VAC is
provided from a stepdown transformer to the navigation unit, the RDDMI and the DAA.
VOR or LOC frequency is selected using the VHF NAV control panel or, when AUTO is enabled,
VOR frequency is automatically provided by the selected FMC through the DAA. AUTO/Manual
selection is made on the VHF NAV control panel.
Antennas used by the VHF NAV unit are selected by the ILS relays. Selected course and
selected runway heading are selected using the Digital Flight Control System (DFCS) Mode
Control Panel (MCP). Magnetic heading to the VHF NAV unit is provided by a bootstrap servo
in the RDDMI.
VOR Operation
VOR signals are received by the VOR/LOC antenna located in the tip of the vertical
stabilizer. The VOR signal is processed by the receiver, and the magnetic bearing to
the VOR station is computed. The VOR bearing is provided through the DAA to the FMCS
and the EFIS symbol generators. The bearing signal is also combined with airplane
heading in the VHF NAV unit to generate a relative bearing output to the RDDMI's via the
integrated flight systems accessory unit. When the FMC selects agility mode for the
DME, a relay in the integrated flight systems accessory unit removes the VOR pointer
information from the RDDMI. The VHF NAV unit provides VOR deviation to FCC A and FCC B.
The NAV unit provides the VOR super flag (28 volt valid signal) to all equipment using
the VOR signals. (The VOR flag signal is no longer used).
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MAINTENANCE TRAINING MANUAL
OPERATION
Localizer Operation
LOC signals are received by the VOR/LOC antenna located in the tip of the vertical
stabilizer or, if approach mode is selected for the DFCS, by the LOC antenna located in
the radome. The localizer signal is processed by the NAV unit and localizer deviation
is computed. Localizer deviation is provided to both FCC's, the Ground Proximity
Warning Computer (GPWC), the standby horizon indicator and the DAA. The DAA sends the
LOC deviation to the EFIS symbol generators and the FMCS.
When a localizer frequency is selected, ILS 28 volts dc is provided to the antenna
switching relays, the DAA, the FCC and the EFIS control panel.
The NAV unit provides the LOC super flag to all equipment using LOC signals.
Glide Slope Operation
The glide slope receiver is enabled when a localizer frequency is selected. The glide
slope signal is received by a glide slope antenna located in the radome. The signal is
processed by the glide slope receiver section of the VHF NAV unit and glide slope
deviation is computed.
Glide slope deviation is provided to the FCC's, the GPWC, the standby horizon indicator,
and to the symbol generators via the VHF NAV transfer relays.
The NAV unit provides the glide slope super flag to all equipment using glide slope
signals.
VHF Navigation Self-Test
Navigation system self-test is enabled through switches on the control panel, through
switches on the front of the NAV unit, or by the DFCS during VHF NAV interface test of
the DFCS BITE operation. Switches provided on the control panel and on the front of the
NAV unit are: VOR, ILS UP/L (up/left) and ILS DN/R (down/right). The DFCS test is ILS
UP/L only.
A VOR frequency must be selected to arm the VOR test and a LOC frequency must be
selected to arm ILS tests. When the approach mode is selected for the DFCS, all
navigation self-tests are inhibited.
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VOR/ILS NAVIGATION SYSTEM BLOCK DIAGRAM
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MAINTENANCE TRAINING MANUAL
OPERATION
GLIDE SLOPE DEVIATION AND GLIDE SLOPE SUPERFLAG INTERFACE
Operation
Glide slope deviation and GLIDE SLOPE SUPERFLAG outputs from the VHF NAV units are sent to
the Flight Control Computers (FCCs), the Ground Proximity Warning Computer (GPWC), and the
Standby Horizon Indicator. The outputs also go to the EFIS Symbol Generators via the VHF
NAV TRANSFER relays. Normally VHF NAV unit-1 supplies the left side and VHF NAV unit-2
supplies the right.
The glide slope scale and pointer will be displayed if a localizer frequency is selected.
If back course is detected, or if glide slope SUPERFLAG is invalid, or if glide slope is
NCD, the pointers are removed. If a glide slope deviation warning is detected in the FCC,
the EADI scale changes to yellow and the pointer flashes at a 4 Hz rate.
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MAINTENANCE TRAINING MANUAL
GLIDE SLOPE DEVIATION AND GLIDE SLOPE SUPER FLAG INTERFACE
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MAINTENANCE TRAINING MANUAL
OPERATION
VHF NAV INTERFACE SCHEMATIC - VOR/LOC
VHF NAV Outputs
VOR/LOC deviation, VOR bearing, VOR/LOC super flag, and ILS 28v dc are sent from the VHF NAV
unit to the DAAs. Within the DAA's these signals are conditioned, multiplexed and converted
to the ARINC 429 format for transmission via an ARINC 429 DATA bus to the FMCS, EFIS Symbol
Generators, and the GPWC.
The ILS 28v dc is also sent to the Standby Horizon indicator, the FCC, and through the NAV
transfer relays to both EFIS control panels. VOR/LOC deviation is also sent to the Standby
Horizon indicator, and both FCC's. VOR/LOC superflag is also sent to the Standby Horizon
indicator, the FCC, and through the agility switch to the RDDMI.
EFIS Signal Processing
Data from the DAA comes into the symbol generator on an ARINC 429 data bus through a switch
controlled by the VHF NAV transfer relay.
Selected course or selected runway heading is supplied to the symbol generators by the FCCs.
If the selected course or selected runway heading data is invalid from the FCCs it is
provided directly from the DFCS Mode Control Panel. This positions the selected course or
selected runway heading pointer on the EHSI and provides selected runway heading to the back
course detector.
ILS
When the VHF NAV system is tuned to a ILS frequency the localizer deviation circuits
within the EFIS symbol generator provide localizer deviation signals to the EADI and
EHSI. The expanded scale detector monitors the deviation signal and causes the
localizer scale on the EADI to expand when VOR/LOC mode is selected on the DFCS mode
control panel and the localizer deviation is less than 5/8 of a dot. The back course
detector monitors the selected runway heading and the airplane track. When the
difference between the two is 90ø or more then the back course signals go out to the
EADI localizer circuits to reverse polarity of the localizer deviation, and to the glide
slope circuits to inhibit the glide slope pointer.
VOR
When the VHF NAV system is tuned to a VOR frequency the VOR deviation circuits within
the EFIS symbol generator compute VOR deviation using VOR bearing, and selected course.
The TO/FROM circuit uses the same inputs to compute TO/FROM. The VOR vector circuits
use VOR bearing to display VOR vectors if selected.
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MAINTENANCE TRAINING MANUAL
VHF NAV INTERFACE SCHEMATIC - VOR/LOC
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MAINTENANCE TRAINING MANUAL
OPERATION
VOR/ILS NAVIGATION SYSTEM SELF-TEST INTERFACE SCHEMATIC
Operation
Self-test of VOR/ILS NAV systems 1 and 2 are similar. Only system 1 is discussed below.
VOR Test
The VOR test signal from the VHF NAV-1 Control Panel is sent through the normally
closed switch U1 (Test Inhibit Switch). During the approach mode or VOR/LOC mode with
an ILS frequency tuned, FCC-A opens switch U1 to interrupt the VOR test input to the
VHF NAV Unit.
ILS TEST
DN/RT and UP/LT ILS test signals from the VHF NAV control panel are sent through the
normally closed switches U3, and U4 respectively (VOR/ILS-1 test inhibit). During
approach mode or VOR/LOC mode with an ILS frequency tuned, FCC-A opens these switches
to interrupt ILS test inputs to the VHF NAV unit.
FMC/CDU DISPLAYS
When an ILS or VOR test is performed, the FMCS SENSOR STATUS page displays "VHF NAV
TEST" while the appropriate test switch is actuated. At the conclusion of the test,
"VHF NAV OK" is displayed. If the VHF system is faulty, the display reads "VHF NAV
FAIL".
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VOR/ILS NAV SYSTEM SELF-TEST
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MAINTENANCE TRAINING MANUAL
SYSTEM TEST
VOR SELF-TEST
Test Preparations
The VOR test preparation and procedure described here is for VOR/ILS navigation system No.
1. The test of system No. 2 is performed in the same way, using the controls and indicators
for system No. 2.
Prior to testing system No. 1, the following switches must be in the position indicated:
VHF NAV Transfer Switch .................................. NORMAL
EFIS Control Panel ....................................... VOR/ILS
Course Select No. 1 ...................................... 000
Next, select a VOR frequency of 108.00 MHz on the VHF NAV Control Panel. Observe the VHF NAV
status on the sensor status display on the flight management computer control display unit
(FMC CDU).
To initiate a VOR self-test, press and hold the VOR TEST switch, and observe the following
displays:
Captain's EHSI
Deviation Bar ............................................
To/From Annunciation .....................................
Mode.....................................................
Course Arrow .............................................
Centered
FROM
VOR 1
000(N)
RDDMIs
No. 1 Pointer (VOR) ...................................... 180° (±3°)
No. 1 Pointer Flags ...................................... *[1]
FMC CDU
VHF NAV Status ........................................... TEST
*[1] RDDMI warning flags remain in view for 3 ñ1 seconds; retract from view for 24 ±6
seconds; reappear 27 ±7 seconds after initiating test for as long as self-test is
actuated. During the time that RDDMI flags are in view, the VOR Deviation Bar is
removed and the TO/FROM annunciation is removed on the EHSI display.
Release the TEST switch and all VOR indications return to normal.
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VOR SELF-TEST
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MAINTENANCE TRAINING MANUAL
SYSTEM TEST
ILS SELF-TEST
General
The ILS up/left and down/right self-test preparation and procedure described here is for the
VOR/ILS navigation system No. 1. The test of system No. 2 is performed in a similar
fashion. Prior to the testing system, the following switches must be in the indicated
positions:
VHF NAV Transfer Switch .................................. NORMAL
EFIS Control Panel ....................................... VOR/ILS
Select a localizer frequency on the VHF NAV control panel. Observe the VHF NAV status on the
sensor status display on the FMC CDU. To initiate the ILS up/left self-test, place and hold
the TEST switch in the UP/LT position and observe the following displays:
Captain's EHSI
Localizer Deviation ...................................... Left 1 Dot
G/S Deviation ............................................ Up 1 Dot
Warning Flags ............................................ *[1]
RDDMIs
Pointer 1 ................................................ 3:00 O'clock Position
Captain's EADI
Localizer Deviation ...................................... Left 1 Dot
G/S Deviation ............................................ Up 1 Dot
Warning Flags ............................................ *[1]
FMC CDU
VHF NAV Status ........................................... TEST
*[1]
RDDMI warning flags remain in view for 3 ±1 seconds; retract from view for 24
seconds; reappear 27 ±7 seconds after initiating test for as long as self-test is
actuated. During the time that RDDMI flags are in view, the LOC deviation
stem/bar and the G/S deviation pointers are removed from the EADI and EHSI.
Release the TEST switch and all ILS indications return to normal. Perform the test again
with the switch in the DN/RT position. The indications are similar to UP/LT test, but this
time the deviation bar and runway symbol is positioned right 1 dot, and the glide slope
deviation pointer is positioned down 1 dot.
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ILS SELF-TEST
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MAINTENANCE PRACTICES
VHF NAVIGATION SYSTEM - MAINTENANCE SUMMARY
Maintenance Summary
Refer to the Maintenance Manual, Chapter 34, Section 31, for all troubleshooting,
testing and component removal or installation procedures.
Power
Prior to operating, testing or troubleshooting the VHF NAV system, assure that the proper
power is available and all related circuit breakers are set.
Tests
A complete self-test of the system is available at the VHF NAV/DME panel. Consult the
Maintenance Manual for particulars.
Removal/Replacement
The VHF NAV system components contain electrostatic discharge sensitive (ESDS) devices.
To prevent damage to these devices refer to the Maintenance Manual, Chapter 20, Section
40-12, for handling instructions.
POWER
POWER TO VHF NAV, HEADING
AND INSTRUMENTS
SYSTEM
TEST
TEST IS AVAILABLE AT THE
CONTROL
PANEL.
DISPLAYS ON THE
CONTROL PANEL
ARE TESTED THROUGH
THE MASTER
TEST SWITCH.
REMOVAL/
INSTALLATION
SYSTEM COMPONENTS CONTAIN
ESDS DEVICES
118535
VHF NAV SYSTEM – MAINTENANCE SUMMARY
62
Marker Beacon
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
MARKER BEACON OPERATION
1. Purpose
The Marker Beacon system indicates to the flight crew that the airplane is passing directly
over specific geographic points. A ground radio station marks the point.
2. System Description
Marker beacon stations all transmit a 75 MHz RF signal modulated with either a 400, 1300 or
3000 Hz audio. This audio is keyed (dots and/or dashes) to provide identification of the
station and illumination of the appropriate flight compartment lights.
Markers are generally used on final approach to the runway. The outer marker is located
approximately four miles from the runway end. When the airplane passes over this marker,
the blue light on the forward instrument panels turns on and a tone (400 Hz) keyed as
continuous dashes is heard on the interphone. The middle marker turns on the amber light on
the panels and a 1300 Hz tone is keyed as alternate dots and dashes. This marker is located
approximately five tenths of a mile from the runway end. The inner marker, located
approximately one tenth of a mile from the runway end, turns on the white light on the
panels and a 3000 Hz tone is keyed as continuous dots. The back course marker located at
the opposite end of the runway from the inner marker also has a 3000 Hz tone and turns on
the white light but the tone is keyed as continuous paired dots. By monitoring the lights
and tone the flight crew is able to mark progress on final approach to the runway.
When the markers are used as airways identifiers, the 3000 Hz tone is used and the white
light on the panels turns on when the airplane passes over the beacon. The tone is keyed
with a Morse code identifier. The specific code is obtained from the airway charts. The
keyed codes for final approach markers is as stated previously unless otherwise altered on
airport area charts.
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MARKER BEACON
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
MARKER BEACON COMPONENT LOCATIONS
1. General Component Locations
A.
Circuit Breaker
The circuit breaker is located on the captain's load control panel (P18-1). This
item will not be discussed further.
B.
Receiver
The marker beacon receiver is installed on the E2-4 shelf in the electrical and
electronic equipment compartment.
C.
Marker Beacon Lights
The light assemblies are installed on the pilots' instrument panels, one each on the P1 and P-3.
D.
Antenna
The antenna is flush mounted on the underside of the airplane, along the centerline at
STA 620.
2. System Interfaces
The marker beacon receiver provides an audio output to the audio integration system.
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MAINTENANCE TRAINING MANUAL
MARKER BEACON COMPONENT LOCATION
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
MARKER BEACON RECEIVER
1. Purpose
The marker beacon receiver processes signals from a ground station and provides a visual
signal in the form of a light and an aural signal (tone) to indicate passage over the ground
station.
2. Features
The unit is 1/4-ATR short/low and weighs 1.47 Kg (3.25 lb) and requires 28 volt dc power.
The receiver frequency is 75 MHz single channel, fixed tuned.
3. Self Test
An optional test switch is mounted on the front panel of some receivers with certain
specific part numbers. When pushed, a sequential test is started which results in lights
and tones.
MARKER BEACON RECEIVER
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
MARKER BEACON LIGHTS
1. Purpose
The marker beacon lights provide a visual indication to the flight crew informing them of
both beacon passage and type.
2. Features
Each marker beacon light assembly consists of three single bulb modules. The top module is
white and is labelled AIRWAYS. The center module is amber and is labelled MIDDLE. The
bottom module is blue and is labelled OUTER.
3. Power
The lights receive dc power from the marker beacon receiver
4. Lamp Test
The lights are tested by either of two procedures. The whole assembly is tested by using
the master dim and test relay. Each module has a push-to-test feature for
individually checking the bulbs.
MARKER BEACON LIGHTS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
MARKER BEACON ANTENNA
1. Purpose
The antenna provides the means of receiving the RF energy from the marker ground station.
2. Features
The antenna is a flush mounted-type antenna, tuned for a 75 MHz signal. The antenna is
installed at STA 620 on the bottom of the airplane and is held in place by six mounting
screws.
MARKER BEACON ANTENNA
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MAINTENANCE TRAINING MANUAL
OPERATION
MARKER BEACON SYSTEM BLOCK DIAGRAM
Operation Sequence
Power
The marker beacon system receives 28 volts dc from the circuit breaker panel and
operates whenever power is available.
Signal Processing
During passage over a ground station, the type of the station is identified by the
pitch and key modulation of the transmitted signal. The receiver processes the received
signal, discriminates the pitch to illuminate the appropriate light and provides audio
signals to the audio selector panels.
The high/low sensitivity of the receiver is automatically controlled by the low range
radio altimeter 1500 foot trip.
MARKER BEACON SYSTEM BLOCK DIAGRAM
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
MARKER BEACON SYSTEM MAINTENANCE SUMMARY
1. Servicing
Refer to Maintenance Manual, Chapter 34, Section 35, for all maintenance procedures.
2. Functional Test
The test switch on the front of the marker beacon receiver enables a system test simulating
a receiver input. The test verifies system operation, excluding the antenna and antenna
interface.
3. Power Sources
Power for the marker beacon system is controlled by a single circuit breaker on the P18 load
control center. The marker beacon system interfaces with the audio integration system which
must also have power applied.
SERVICING
MARKER BEACON SYSTEM MAINTENANCE IS COVERED IN THE MAINTENANCE MANUAL UNDER
CHAPTER 34, SECTION 35.
FUNCTIONAL
TEST
MARKER BEACON TEST CHECKS SYSTEM FROM RECEIVER INPUT THROUGH THE MARKER
BEACON LIGHTS.
POWER SOURCES
ASSURE THAT POWER IS APPLIED TO THE MARKER BEACON SYSTEM.
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MARKER BEACON SYSTEM MAINTENANCE SUMMARY
10
Low Range Radio
Altimeter
2
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MAINTENANCE TRAINING MANUAL
NOTES
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
LOW RANGE ALTIMETER SYSTEM
1. Purpose
The low range radio altimeter (LRRA) system measures the vertical distance (absolute
altitude), from the airplane to the terrain.
2. System Description
The low range radio altimeter system operates from -20 to 2500 feet and is used primarily
during the approach and landing phases of the flight.
The LRRA system measures altitude by transmitting a signal to the ground and comparing this
signal to the signal reflected from the terrain. The operating center frequency is 4300
MHz.
The R/T compares the reflected and transmitted signals and converts the difference to a
signal that drives the indicator. The R/T also supplies altitude information to other
airplane systems.
LOW RANGE RADIO ALTIMETER SYSTEM
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
LOW RANGE RADIO ALTIMERER SYSTEM COMPONENT LOCATION
General Component Location
Circuit Breaker
The circuit breaker for LRRA No. 1 R/T and No. 2 R/T are located on the P18-1 and P6-1
panels, respectively.
Transceivers
Two receiver/transmitters are located on the E2-4 shelf of the electrical/electronics
compartment.
Antennas
Two transit and two receive antennas are located on the bottom centerline of the
airplane.
Indicators
One radio altimeter indicator and one attitude director indicator with a rising runway
and decision height (DH) light, located on each of the captain’s and first officer’s
instrument panels.
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MAINTENANCE TRAINING MANUAL
LOW RANGE RADIO ALTIMETER SYSTEM COMPONENT LOCATIONS
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
LRRA SYSTEM RECEIVER/TRANSMITTER
1. Purpose
The LRRA R/T unit provides the required signal generation, reception and processing
necessary to measure the absolute height of the airplane above the ground.
2. Physical Description
The radio altimeter is contained in a 1/2 ATR short case; weighs 15 pounds. The unit is
secured to the shelf by two hold down fasteners. A handle is provided to facilitate
carrying the unit. The front panel test connector is used for flight line testing using a
radio altimeter test set.
3. Power
The transceiver requires 115 volts ac, 400 Hz, single-phase electrical power. All electrical
connections are made through a single rear connector.
4. Operation
The transmitter produces a frequency modulated continuous wave (FMCW) output centered at
4300 MHz. The receiver has subassemblies which receive the ground-reflected signals,
compares them to the transmitter frequency and produces signals that are proportional to
absolute altitude. The altitude signal processing is performed by two microprocessor units
within the R/T unit. One microprocessor performs the altitude processing which produces
analog and digital output signals; the other processor performs monitoring functions. The
R/T unit also provides altitude trip capabilities that can be used by autopilot and/or
flight control systems.
5. Self-Test
Test switch/status indicator; during normal operation the switch/indicator is not
illuminated. When pressed, the switch/indicator activates a self-test of the unit. A
successful test illuminates the switch/indicator and causes the LRRA indicator to read 40
feet.
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MAINTENANCE TRAINING MANUAL
LRRA SYSTEM RECEIVER/TRANSMITTER
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MAINTENANCE TRAINING MANUAL
NOTES
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
LRRA SYSTEM ANTENNA
1. Purpose
The antenna provides the means for radiating or receiving RF energy.
2. Features
The antenna is connected to the R/T by a coax cable whose length is critical. Each antenna
is held in place by 8 screws and has a keying pin to insure correct orientation. Transmit
and receive antennas are indentical for each system and are interchangeable.
3. Operation
The antenna operates in the 4300 MHz frequency range.
LRRA SYSTEM ANTENNA
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
LRRA SYSTEM INDICATOR
1. Purpose
The LRRA indicator provides an analog display of radio altitude from 0 to 2500 feet. The
indicator shows system status and provides means to set reference altitude and to test the
system.
2. Features
The indicator scale, which may be illuminated, is linear from 0 to 500 feet and logarithmic
from 500 to 2500 feet. A warning flag in view signifies that the altitude displayed is
invalid. A decision height (DH) index is positioned by the rotation of the SET/TEST
selector knob. The decision height is also displayed digitally on the DH counter. Pressing
the SET/TEST selector knob actuates complete test of the indicator and receiver/transmitter.
3. Power
Power requirements are 115 volts ac, 400 Hz and ±30 volts dc. The indicator's panel lights
require 5 volts ac, 400 Hz.
4. Operation
The output signal from the LRRA receiver/transmitter drives the altitude tape behind the
reference mark. If the measured altitude exceeds 2500 feet, the airplane reference symbol
points to the black section of tape above the 2500 feet mark on the altitude tape.
The SET/TEST knob, when rotated, sets the decision height (DH) counter and moves the DH tape
to align the DH index to the selected altitude on the display tape. The fixed DH reference
symbol provides a reference point for the DH index on the DH tape to indicate the airplane
is above or below the DH.
The DH amber indicator illuminates when the actual altitude is equal to or less than the DH
altitude. The DH light can subsequently be extinguished by pressing the light.
5. Test
Successful test results are an altitude display of 40 feet concurrent with the flag in view.
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MAINTENANCE TRAINING MANUAL
LRRA SYSTEM INDICATOR
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
ADI RISING RUNWAY SYMBOL
1. Purpose
The rising runway symbol on the attitude director indicator (ADI) provides the pilot with a
two-dimensional representation of the airplane's position relative to the runway centerline
during the approach phase of flight.
2. Features
The horizontal displacement of the symbol is a function of localizer deviation; the vertical
displacement is a function of altitude measured by the LRRA.
3. Operation
The runway symbol is visible only when the VHF navigation system is tuned to an ILS
frequency. When operating on localizer at altitudes above 200 feet, the symbol displays
localizer steering information only and appears just above the localizer scale. When the
airplane descends below 200 feet on its approach, the runway symbol rises vertically
proportional to the airplane's altitude above the runway. At touchdown, the runway symbol
should be at the lower edge of the fixed airplane symbol.
4. Monitor
Circuitry in the ADI monitors the validity of the LRRA R/T and the runway symbol operation.
Malfunction of either system causes the RUNWAY flag to be deployed.
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MAINTENANCE TRAINING MANUAL
ADI RISING RUNWAY SYMBOL
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
LRRA SYSTEM THEORY
Operation Sequence
Time/Frequency Relationship
The system transmits an FM signal of linearly varying frequency which, when delayed by
the round-trip time to the ground and then mixed with a portion of the transmitted
signal, gives a difference frequency that is proportional to the airplane height above
the ground.
Operation
The transmitted frequency range is 4300 ±A MHz. At some point in time (t1), the
transmitted frequency (f1) is transmitted to the ground and the reflected signal
received at the LRRA's mixer at time (t2). During the time interval (∆t), the
transmitted frequency has increased to a new frequency (f2). When the frequencies f1
and f2 are mixed at time (t2), the output of the mixer is the difference frequency (∆f)
which is proportional to ∆t and therefore to the altitude (H). By means of proper
calibration, this difference in frequency is converted to a dc voltage and sent to
interfacing systems.
LRRA SYSTEM THEORY
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MAINTENANCE TRAINING MANUAL
OPERATION
LRRA SYSTEM BLOCK DIAGRAM
Operation Sequence
Power
Each LRRA receiver-transmitter (R/T) receives power from a 115 volt ac source, R/T 1
from the ELEX PWR 1 bus and R/T 2 from the ELEX PWR 2 bus.
Signal Processing
Each R/T transmits signals through its transmitting antenna to ground and the reflected
signals are returned to the receiving antenna. The R/Ts process the RF signals and
produce altitude outputs which are applied to the LRRA indicators, the attitude
director indicators (ADIs), the flight control computers (FCC) and the instrument
warning comparator. The LRRA system No. 1 also supplies the ground proximity warning
computer, the A/T computer and the Yaw Damper Computer. The R/Ts also supply a number
of trip signals which are set at various altitudes. The 500 foot trips enable altitude
comparison in the instrument comparator. The 200 foot trips enable the rising runways
in the ADIs. The 10-foot trips activate the thrust reversers when the thrust reversers
are armed. LRRA indicator 1 supplies DH to the Ground Proximity Warning Computer. The
1500-trip controls the Marker Beacon Receiver sensitivity. Valid signals from the R/Ts
control flag positions in the pilots' and control the warning circuits in the
interfacing avionic systems. The LRRA indicator senses decision height (DH) and
illuminates the DH light on the ADI.
Test
Each system is tested from the associated LRRA indicator. During the test, the
indicator displays a specific altitude reading and the warning flag comes into view.
The test is interlocked through the instrument accessory unit to the R/T. When the
associated autopilot is armed or engaged in the ILS mode, the test is inhibited.
During a test of the LRRA system 1, the ground proximity warnings are inhibited.
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MAINTENANCE PRACTICES
LRRA SYSTEM - FUNCTIONAL TEST
Functional Test
General
The captain's and first officer's LRRA systems are independent of one another and their
self-tests are performed independently.
A self-test of the system can be initiated from the R/T unit or the LRRA indicator.
However, the LRRA indicator TEST switch is inhibited. When the autopilot is in GS
capture mode.
Procedure
Provide electrical power and check that LRRA 1 circuit breaker on panel P18 and LRRA 2
circuit breaker on panel P6 are closed.
Set DH index of the LRRA indicator to zero feet and check that the DH light is
extinguished. The altitude tape on the indicator indicates between 0 and -10 feet and
the warning flag on the indicator is out of view. The DH light is disabled for
altitudes less than 10 feet regardless of the position of the DH index.
Depress SET/TEST button on the LRRA indicator or the test/status switch on the R/T unit
and check that the DH light is extinguished, the altitude tape on the indicator indicates
40 ±5 feet and the warning flag on the indicator is in view.
Slowly adjust DH index on the indicator to increasing altitude setting and check that
the DH light illuminates when DH index setting is within ±5 feet of displayed test
altitude.
Release SET/TEST button and check that captain's DH light remains illuminated.
Set DH index back to zero feet and check that captain's DH light goes out.
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LRRA SYSTEM - FUNCTIONAL TEST
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MAINTENANCE PRACTICES
LRRA SYSTEM MAINTENANCE SUMMARY
Summary
General
Refer to the Maintenance Manual, Chapter 34, Section 33 for all troubleshooting,
component removal and installation procedures.
Since the altitude indicator is not calibrated below zero feet, the altitude displayed
on the ground is an approximation (0 to -10 feet).
Power
There is a separate circuit engaged for each R/T. The radio altitude circuit breakers
must be set to enable the rising runway function on the ADIs.
Tests
System test may be actuated from the flight compartment or from the
electrical/electronics equipment compartment.
Removal/Installation
Improper antenna bonding can cause system malfunctions during single flights that
typically read as "intermittent flags and erratic altitude deviations from 400 feet to
2500 feet". Be certain gasket is in good condition and flying surfaces are clean.
POWER
ASSURE THAT POWER IS APPLIED TO ALL RELATED SYSTEMS.
TEST – LOCAL
(FLIGHT COMPARTMENT)
SYSTEM TEST AT INDICATOR. LRRA INDICATOR POINTER DRIVEN TO PRESET
ALTITUDE AND WARNING FLAG IN VIEW. RISING RUNWAY
ON ADI DRIVEN TO SPECIFIED ALTITUDE. DH/MDA LIGHT ILLUMINATES IF THE TEST ALTITUDE IS BELOW THE DH/MDA SETTING.
TEST - REMOTE
ELECTRICAL/ELECTRONICS
EQUIPMENT COMPARTMENT
SYSTEM TEST AT R/T UNIT. SAME FLIGHT COMPARTMENT INDICATIONS AS IN LOCAL TEST. TEST/STATUS LAMP ILLUMINATES
DURING SUCCESSFUL TEST. DURING NORMAL OPERATION, STATUS
LAMP IS NOT ILLUMINATED.
REMOVAL/INSTALLATION
ASSURE PROPER ANTENNA BONDING
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EFIS Differences
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COMPONENT FUNCTIONAL DESCRIPTION
EFIS CONTROL PANEL
Purpose
The EFIS Control Panel is used to select the desired display on the EADI and EHSI. The
control panel is divided into two halves. The left side controls the EADI and the right
side controls the EHSI. The following text is confined to the controls for the EADI.
Features
Decision Height Select Knob
This 24-detent, continuous-turn control knob adjusts the decision height between -20
feet to +999 feet. Normally, rotating the knob one detent changes the DH setting one
foot; however, rotating the knob faster than two revolutions/second changes the DH
setting by four feet/detent.
DH REF LCD Display
This display shows the selected decision height. On power-up, the display shows 200
feet. During a master light test, the display alternates between "888" for 2 seconds
and blank for 1 second.
RST
After the airplane has descended through the decision height, the DH circuits can be
manually reset by pressing the RST pushbutton switch. The RST pushbutton can also be
used to reset the height alert.
BRT
This knob adjust the brightness of the EADI display.
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EFIS CONTROL PANEL
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NOTES
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OPERATION
RADIO ALTIMETER SYSTEM - EADI DISPLAY
Purpose
The EADI displays the radio altitude and decision height for use primarily during approach
and landing.
Features
Radio Altitude Display
Radio altitude is displayed for altitudes between -20 and 2500 feet. From -20 to 100
feet the display updates in 2-foot increments; between 100 and 500 feet the display
updates in 10-foot increments, and from 500 to 2500 feet the display updates in 20-foot
increments. Above 2500 feet the display is blank. If the radio altitude is greater
than 1000 feet for greater than 2 seconds a digital readout is displayed. If the radio
altitude is less than or equal to 1000 feet for greater than 2 seconds a radio altitude
(RA) dial with readout is displayed. The complete RA dial shall represent an RA of 1000
feet.
Rising Runway
The green rising runway will be displayed when LOC is valid. For altitudes from 2500
to 200 feet the runway remains at 200 feet. At 200 feet and below the runway will rise
with decreasing airplane altitude until it meets the airplane symbol.
Decision Height Display
The decision height displayed on the EFIS Control Panel is displayed above the radio
altitude for radio altitudes greater than 1000 feet. For radio altitudes less than or
equal to 1000 feet the DH index is displayed on the RA dial. If the DH selected is
negative, the digital decision height or DH index is blanked.
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OPERATION
Operation
No DH Alert (RA greater than DH)
The DH selected is displayed above the radio altitude display for radio altitudes greater
than 1000 feet. If the radio altitude is less than or equal to 1000 feet the RA dial and DH
index with the standard readout is displayed.
DH Alert (RA less than or equal to DH)
As the airplane descends through the selected decision height value, the entire display –
radio altitude dial indices, ring segment and DH index - change to yellow and flash during
the first three seconds. A DH-Alert is possible only if the airplane descends from a
radio altitude value equal to or greater than selected decision height value plus 75 feet.
Decision Height Alert Termination
The decision height alert can be terminated automatically or by manual reset. Automatic
reset occurs at touchdown, or when the airplane climbs to a height greater than 75 feet
above the selected decision height. Manual reset is achieved by actuating the reset
pushbutton switch RST on the EFIS control panel.
At reset, the display returns to the normal colors: the radio altitude changes back to
white, the dial indices and ring segment change back to white and the DH index changes back
to magenta.
Invalid Data
Invalid DH or radio altitude are indicated by yellow displays of a boxed "DH" or a boxed
"RA", respectively. Invalid radio altitude also blanks the rising runway display
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RADIO ALTIMETER SYSTEM - EADI DISPLAY
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OPERATION
LRRA SYSTEM BLOCK DIAGRAM
Operation Sequence
Power
LRRA systems 1 and 2 receive power from circuit breaker connected to the 115 volt ac
ELEX Power Busses 1 and 2, on the P18 and P6 panels, respectively.
Signal Processing
The LRRA transmits a signal from the transmit antenna and the reflected signal is
returned to the receive antenna. In the R/T, the receive and transmit signals are
mixed together to produce an altitude output which is connected to the EFIS symbol
generators, the ground proximity computer, the autothrottle (A/T) computer, the flight
control computers, TCAS and the yaw damper computer. The digital flight data recorder
system receives the altitude through the symbol generators.
The LRRA R/T contains detectors that monitor the altitude and provide trip signals at
certain altitudes. The 10-foot trips enable the thrust reversers. The 1500' trips
cause the marker beacon receiver to switch to low sensitivity below 1500'.
The LRRA R/Ts contain monitor circuits which supply VALID signals when the R/Ts are
operating normally.
A decision height (DH) is selected on the EFIS control panel. When the airplane decends
to or below the selected DH, the DH alert is displayed on the EADI. When the airplane
ascends 75 feet above the DH level or the reset switch is pressed, the DH alert display
extinguished.
Test
The DH alert can be tested through the EFIS test. The EFIS test is performed through
the EFIS bite.
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LRRA SYSTEM BLOCK DIAGRAM
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SYSTEM TEST
OPERATIONAL TEST
LRRA Self-Test
The LRRA may be tested by pushing and holding the self-test switch on the front of the
receiver/transmitter. When the self-test is activated, the receiver/transmitter sends both
an altitude of 40 feet and an RA invalid signal to all systems. On the EADI when the test
switch is pushed, the radio altitude will increase and may display 40 feet momentarily.
After approximately 2 seconds, the EADI will display an RA flag until the test button is
released. If the self-test is okay, the status light on the receiver/transmitter will
illuminate after a 2-second delay and will remain illuminated until the switch is released.
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LRRA OPERATIONAL TEST
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NOTES
32
Distance Measuring
Equipment
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
DME SYSTEM
1. Purpose
The distance measuring equipment (DME) system provides the flight crew, and associated
systems, with the slant range distance to a selected navigation facility. The system is
designed to indicate a maximum range of 389.99 nautical miles.
2. System Description
The DME system consists of a control panel, an interrogator, an antenna and a distance
display. DME control is provided on the VHF Nav control panel.
The DME system measures the distance to the navigation facility by transmitting a pulse pair
signal to the ground station and measuring the time taken to receive a reply to the
interrogation.
The DME channel is automatically selected when the crew selects the frequency of the VHF
Navigation facility (VOR or Localizer). DME channel selection and self-test are controlled
using the VHF Nav control panel. VHF Nav frequency and DME channel may be automatically
selected by the Flight Management System.
DME distance is displayed by digital indicators located on the pilot's instrument panels.
The distance signal is also provided to the Flight Management System for use by the FMC and
for distribution to other systems.
DME audio is provided to the Audio Integrating System to allow aural identification of the
ground facility.
Suppression signals between the ATC transponders and DME interrogators prevent mutual
interference between the systems.
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DME SYSTEM
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GENERAL DESCRIPTION
DME SYSTEM COMPONENT LOCATION
General Component Locations
Circuit Breakers
There is one ac circuit breaker for each system. The circuit breaker for DME-1 is
located on the P18-1 on the left load control center. The circuit breaker for DME-2 is
located on the P6-1 on the right load control center.
Control Panels
There is one VHF NAV/DME control panel for each DME system. Both are located on the
aft electronics panel (P8); No. 1 is on the left and No. 2 is on the right.
Interrogators
Two interrogators are located on the E2-2 shelf of the electronic equipment
compartment. When facing the rack, interrogator No. 1 is on the left and No. 2 is on
the right.
Antennas
The antennas are located on the underside of the fuselage; No. 1 is at station 468 and
No. 2 is at station 580.
Indicator
The DME distance display is part of the radio direction distance magnetic indicator
(RDDMI). There are 2 RDDMIs; one is located on the P-1 panel and the other is on the
P-3 panel.
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DME SYSTEM COMPONENT LOCATION
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COMPONENT FUNCTIONAL DESCRIPTION
VHF NAV/DME CONTROL PANEL
1. Purpose
The panel provides for the selection of automatic or manual frequency control for the VHF
NAV and DME systems. Controls are provided to select the frequency during manual operation.
Segmented displays are provided for both the manual and automatic frequency selections.
Switches are provided to initiate self-test of the VHF NAV and DME systems.
2. Features
A.
AUTO/MAN Selector Switch
The AUTO/MAN selector switch is a push-on/push-off illuminated switch that is used to
select either AUTO or MAN frequency control of the VHF NAV and DME systems. The
selected frequency control mode is indicated on the switch. The annunciator legends are
white on black background, and are visable only when the internal lamps are illuminated.
AUTO selection is armed by the airplane NAV/VOR-ILS switching. NAV must be selected for
display on the HIS before AUTO frequency control can be selected.
B.
Frequency Selector Controls
Dual concentric controls are provided for the manual frequency selection. The outer knob
controls tens and units, and the inner knob controls tenths and hundredths. The
frequency range is 108.00 to 117.95 MHz, selectable in 50 kHz steps. In AUTO mode, the
frequency range is extended to 135.95 MHz.
C.
Frequency Displays
The MANUAL frequency display shows the manual frequency selected. When the system is
in the AUTO mode, a dayglo fire orange FLAG (bar) is raised over the manual display
window. The frequency selected by the flight management system appears in the AUTO
display or dashes are displayed indicating agility tuning. In the MAN mode, the flag
is retracted from the manual display and the automatic display is blanked. When the
master dim and test switch is actuated, both displays read 188.88, flashing 2 seconds
on and 1 second off. Both the AUTO and MAN annunciations are illuminated steadily.
D.
Test Switches
Two switches (three-position, spring-loaded to center) are provided to self-test the
VHF NAV and DME systems. Activation of a test switch sends a test signal to the
appropriate system to initiate a self-test. Activation of a VOR, UP/LT, or DN/RT test
interrupts the flight management system interlock circuit.
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VHF NAV/DME CONTROL PANEL
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COMPONENT FUNCTIONAL DESCRIPTION
DME INTERROGATOR
1. Purpose
The DME interrogator transmits interrogation pulse pairs to a ground station, receives reply
pulses, processes the signals, and provides slant range distance to the DME display and other
airplane systems.
2. Physical Description
The interrogator is enclosed in a rack mounted 1/2 ATR case and weighs 16 lbs. (7.3 kg). It
is secured by two hold-down devices. Interface with the airplane wiring is through a rearmounted dual connector.
3. Power
The interrogator requires 115 volts ac, 400 Hz.
4. Operation
Each DME channel is paired with a VHF Navigation frequency. DME receive frequencies are in
channels between 962 and 1213 MHz. The transmit frequencies are 63 MHz above or below the
receive frequencies, and are in the band of 1025 to 1150 MHz. There are 252 channels
composed of X and Y channels. Currently, only the X channels are used. The X channel pulse
pair spacing is 12 microseconds and the Y channel spacing is 36 microseconds. The output
power of the transmitter is a nominal 700 watts peak.
The slant range capability of the interrogator is from 0 to 389.99 nautical miles.
The interrogator contains monitor circuitry that continuously monitors system performance.
If an annomaly is detected, indicators and interfacing systems are alerted by loss of a DME
VALID signal.
5. Self-Test
The interrogator contains self-test circuitry that simulates normal interrogation and reply.
Self-test is initiated by an external test ground, and the result is evaluated by observing
the DME distance displays.
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DME INTERROGATOR
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
DME ANTENNA
1. Purpose
The DME antenna provides the means of radiating the RF energy from the DME interrogator's
transmitter, and receiving RF reply signals transmitted by the ground station.
2. Physical Description
The DME antenna is a blade-type antenna, 3 inches long, and 4 inches across at the base.
The DME antenna is designed for L-band use. It has a nominal RF impedance of 50 ohms and a
maximum VSWR of 2:1 over the frequency range.
3. Access
The antenna has an O-ring moisture seal and is attached to the airplane by four screws.
Interface with the antenna is through a coaxial cable and connector that is connected to the
antenna before the antenna is installed on the airplane. The antennas for the DME and ATC
systems are identical and are interchangable.
DME ANTENNA
COMPONENT FUNCTIONAL DESCRIPTION
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MAINTENANCE TRAINING MANUAL
RDDMI DME DISTANCE DISPLAYS
1. Purpose
The DME indicators display slant range distance to the selected VHF navigation facility.
The distance is used to aid navigation.
2. Physical Description
The RDDMI displays the distance computed by the DME interrogators. The RDDMI also provides
the pilot with compass heading and ADF or VOR bearing. Only the DME display functions are
discussed here. The DME indicators are located on the upper portion of the RDDMI. Each
indicator consists of four magnetic wheel indicator segments displaying hundreds, tens,
units, and tenths of miles. Each indicator has a red and white flag that is retracted out
of view during normal DME operation.
3. Operation
The indicators receive binary coded inputs from the DME
display slant range distance in a digital format on the
indicators. When an absence of computed data exists, a
display. A loss of valid input data, or a faulty power
striped flag obscuring the display.
interrogators, decode the data and
two sets of magnetic wheel
series of dashes appears in the
supply, results is the red and white
When the DME interrogator is auto tuned dashes indicate either no computed data or agility
tuning.
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RDDMI DME DISTANCE DISPLAYS
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OPERATION
DME SYSTEM BLOCK DIAGRAM
Operation
There are two DME systems installed. The following description pertains to system 1;
however, the operation of system 2 is identical.
DME-1 receives power from 115 volts ac ELEX Bus 1 and DME-2 from ELEX BUS 2.
When the DME is in an operating mode, a pulsed signal in the 1025-1150 MHz band is
transmitted (frequency is determined by the Channel Select input). The DME ground
station receives the interrogation and transmits a reply to the interrogator. The
reply frequency is either 63 MHz above or below the interrogation frequency. The
interrogator receives the reply, measures the time delay between the interrogation and
reply pulses and converts the delay time to the equivalent DME distance.
DME distance is provided to both RDDMIs and a D/A adapter, using a six-wire ARINC
568 digital data bus. Each RDDMI displays the distances measured by both DME systems. The
DAA converts the ARINC 568 input to an ARINC 429 output and distributes the DME distance
to the FMC and FCC-A. DME audio is provided to the audio integrating system where it is
mixed with VHF Nav audio to provide identification of the ground station.
The DME channel is selected when the VHF Nav frequency is selected. This is accomplished
using the VHF Nav control panel, or when AUTO is selected on the control panel, tuning of
the VHF Nav and DME systems is accomplished by the flight management computer. Nav
data must be selected on NAV - VOR/ILS switch to arm AUTO tuning.
The DME interrogators and ATC transponders are interconnected by a suppression cable.
When one system is transmitting, a suppression pulse is provided to the other three
systems inhibiting their operation to prevent mutual inteference.
A DME interrogator that is capable of agility tuning is identified by a ground to the
FMC. When the FMC detects a malfunction of one DME, it alternately tunes the valid DME
interrogator to both channels.
DME self-test is initiated by a test switch on the VHF Nav control panel. The DME system is
also tested when performing the autopilot BITE test.
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DME SYSTEM BLOCK DIAGRAM
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
DME SYSTEM TESTS
Functional Tests
Two tests related to the DME system can be performed on the airplane.
DME System Self-Test performs an automatic sequence of tests which result in a
specific sequence of responses.
The DME Control Panel Display Test, which is part of the master light test, confirms
the proper operation of the frequency displays and tuning mode annunciation on the
VHF NAV/DME control panel.
The DME system self-test is initiated through the VOR/DME test switch located on the VHF
NAV/DME control panel, or automatically during digital flight control system BIT operation.
The test result is evaluated by observing the distance display, and by observing the
FMCS Sensor Status Display on the FMC control display unit (FMC-CDU).
When the VOR/DME test switch is placed in the DME position, the following sequence of
displays indicate a valid system.
For two seconds, the DME distance display is obscured by the red and white striped
flag and DME...FAIL is displayed on the FMC-CDU.
For two seconds, the DME distance display indicates dashes, and DME...TEST is
displayed on the FMC-CDU.
Until the switch is released, the test distance of 000.0 NM is displayed, and
DME...TEST continues to be displayed on the FMC-CDU.
When the switch is released, the distance display continues to display 000.0 NM
for 13.5 seconds, and then returns to the pre-test condition. DME...OK is
displayed on the FMC-CDU immediately upon switch release.
The display test is initiated by placing the lighting master dim and test switch, located
on the P2 panel, in the TEST position.
Proper operation of the frequency displays and the AUTO/MAN annunciator is confirmed by
observing the following sequence.
The AUTO and MANUAL frequency displays alternately display 188.88 (2 seconds) and
blank (1 second) until the switch is released.
The AUTO and MAN annunciations illuminate steadily until the switch is released.
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DME SYSTEM TESTS
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NOTES
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EFIS Differences
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COMPONENT FUNCITONAL DESCRIPTION
EHSI DME DISTANCE DISPLAYS
Purpose
The EHSI displays the distance computed by the DME interrogators. The distance is used to
aid navigation.
Physical Description
Only the DME display functions are discussed here. The distance is displayed in the upper
left corner. The 3-digit display shows miles above 100 nautical miles and tenths of miles
below 100 nautical miles.
Operation
The distance is displayed in the VOR and ILS modes when the VHF navigation system is
manually tuned. If there is no computed data, dashes replace the distance. If there is a
DME failure, the complete DME display blanks.
With AUTO selected on the VHF/NAV control panel, the DME distance will not be displayed on
the EHSI. The frequency display in the lower right corner will be replaced with the word
AUTO.
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EHSI DME DISTANCE DISPLAYS
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NOTES
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OPERATION
DME SYSTEM BLOCK DIAGRAM
Power
The DME interrogators receive 115 volts ac from the elex busses.
Frequency Selection
The DME system can be tuned by the VHF NAV/DME control panel. The DME frequency is paired
to a VHF NAV frequency. When a VHF NAV frequency is selected on the control panel, the DME
interrogator tunes to the paired DME frequency.
The DME system can also be tuned automatically by the Flight Management Computer System
(FMCS). When AUTO is selected on the control panel, the VHF NAV and DME systems are tuned
by the selected flight management computer via the digital analog adapter (DAA).
A DME interrogator that is capable of agility tuning provides a ground to the FMC. When the
FMC commands agility tuning, it will tune the interrogator alternately to two different
frequencies as long as AUTO tuning is selected for the interrogator.
Interrogation
When the DME system is in an operating mode, the interrogator transmits an interrogation
pulse pair to the ground station. The ground station receives the interrogation and
transmits a reply. The DME interrogator receives the reply from the ground station and
processes the signal to determine slant range.
The DME interrogators and ATC transponders are interconnected by a suppression cable. When
one system is transmitting, a suppression pulse is provided to the other systems inhibiting
their operation to prevent mutual interference.
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MAINTENANCE TRAINING MANUAL
OPERATION
Distance Signals
The distance data from the DME interrogator is provided to both RDDMIs and the DAA on a sixwire ARINC 568 digital data bus. Each of the RDDMIs display DME distance from both DME
systems. The DAA converts the ARINC 568 input to an ARINC 429 output and sends DME distance
to both FMCs, the FCC, and EFIS. The captain's EHSI normally displays DME distance from the
number one DME system, and the first officer's EHSI normally displays DME distance from the
number two DME system. If the VHF NAV transfer switch is set to both on one or both on two,
then both EHSI's will display DME distance from the selected source.
The DME interrogator sends DME audio identification signals to the Digital Audio Control
System.
Testing
Self-test of the DME system can be initiated by the DME test switch of the VHF NAV/DME
control panel. The DME system is also tested during DFCS BITE.
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MAINTENANCE PRACTICES
DME SYSTEM TESTS
Functional Tests
Two tests related to the DME system can be performed on the airplane.
DME System Self-Test performs an automatic sequence of tests which result in a
specific sequence of responses.
The DME Control Panel Display Test, which is part of the master light test,
confirms the proper operation of the frequency displays and tuning mode
annunciation on the VHF NAV/DME control panel.
The DME system self-test is initiated through the VOR/DME test switch located on the VHF
NAV/DME control panel, or automatically during digital flight control system BITE
operation.
The test result is evaluated by observing the DME distance display. The FMCS Sensor
Status Display on the FMC control display unit (FMC-CDU) will display DME...TEST during the
DME self-test.
When the VOR/DME test switch is placed in the DME position, the following sequence of
displays indicate a valid system.
For two seconds, the DME distance display is obscured by the red and white striped
flag on the RDDMI and the distance is blank on the EHSI.
For two seconds, the DME distance display indicates dashes on the RDDMI and dashes
are displayed on the EHSI.
Until the switch is released, the test distance of 000.0 NM is displayed on the
RDDMI and 00.0 is displayed on the EHSI. When the switch is released, the
distance display continues to display 000.0 NM for 13.5 seconds, and then returns
to the pre-test condition.
The display test is initiated by placing the lighting master dim and test switch,
located on the P2 panel, in the TEST position.
Proper operation of the frequency displays and the AUTO/MAN annunciator is confirmed by
observing the following sequence.
The AUTO and MANUAL frequency displays alternately display 188.88 (2 seconds) and
blank (1 second) until the switch is released.
The AUTO and MAN annunciations illuminate steadily until the switch is released.
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DME SYSTEM TESTS
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
DME SYSTEM MAINTENANCE SUMMARY
Operational Checkout
General
Refer to Maintenance Manual, Chapter 34, Section 55, when performing maintenance on the
DME systems.
Power
Systems requiring power for DME system operation/test are: DME and VHF Nav. The DME
operational test is normally performed in MAN; however, the FMCS must be functional if the
DME is operated in AUTO tuning test mode.
Test
Operational test uses the DME self-test feature. A complete system test, using a DME
transponder test set, is described in the Maintenance Manual.
Distance Display
Assure that both distance displays provide the same indications. Each DME system
drives the two indicators in parallel.
AUTO/MAN
The operational test is normally performed in the manual tuning mode. If the AUTO mode is
selected, and the FMCS has not been initiated, the automatic tuning lines are open and
the AUTO frequency display is blanked.
Antenna
The DME and ATC antennas are both L-Band antennas and have the same part number.
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MAINTENANCE TRAINING MANUAL
PILOT
DETERMINE MODE WHEN MALFUNCTION OCCURRED
POWER
ASSURE THAT POWER IS APPLIED TO ALL RELATED SYSTEMS
TEST
NO TEST EQUIPMENT REQUIRED FOR OPERATIONAL TEST.
DISTANCE DISPLAY
DME DISTANCE IS DISPLAYED ON BOTH RDDMI:
AUTO/MAN
OPERATIONAL TEST IS PERFORMED IN MAN
ANTENNA
DME AND ATC ANTENNAS ARE THE SAME TYPE
DME SYSTEM MAINTENANCE SUMMARY
29
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NOTES
30
Ground Proximity
Warning System
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NOTES
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
GROUND PROXIMITY WARNING
1. Purpose
The purpose of the Sundstrand-MARK ground proximity warning system (GPWS) is to alert the
flight crew to the existence of an unsafe condition due to terrain proximity.
2. System Description
A.
GPWS Modes
The various hazardous conditions that can be encountered in flight are divided into the
following modes:
Mode 1 - Excessive descent rate.
Mode 2 - Excessive closure rate with respect to rising terrain.
Mode 3 - Excessive altitude loss during climb-out (in takeoff or during goaround) when not in landing configuration (landing gear up and/or flaps
less or equal to 15 units).
Mode 4 - Insufficient terrain clearance when not in landing configuration
(landing gear up and/or flaps less than or equal to 15 units).
Mode 5 - Excessive deviation below glide slope when making a front course approach
with the gear down.
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INTRODUCTION
B.
Mode Annunciations
The GPWS modes are annunciated in the flight compartment by means of aural messages and
lights.
(1)
Aural Messages
The aural messages are transmitted over the audio integration loudspeaker.
(2)
Lights
Warning annunciations include the illumination of the red PULL UP lights.
Alert annunciations include the illumination of the amber BELOW G/S P-INHIBIT
lights.
One other light (amber) on the first officer's instrument panel, INOP, will
annunciate an inoperative system. There is no aural message associated with
this annunciation.
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MAINTENANCE TRAINING MANUAL
MODE 1 - EXCESSIVE DESCENT RATE
MODE 2 - EXCESSIVE CLOSURE RATE
MODE 3 - EXCESSIVE ALTITUDE LOSS DURING CLIMB-OUT
MODE 4 - INSUFFICIENT TERRAIN CLEARANCE
MODE 5 - EXCESSIVE DEVIATION BELOW GLIDE SLOPE
GROUND PROXIMITY WARNING SYSTEM
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
GROUND PROXIMITY WARNING SYSTEM
General Component Location
Circuit Breakers
The GPWS circuit breaker is located on the P18 panel.
Control Panels
There is one control panel for the system. It is located on the first officer's P3
panel.
Computer
The GPWS computer is located on the E2-4 shelf.
Warning Lights/Inhibit Switches
A PULL UP warning light and BELOW GS warning light/inhibit switch are located on each
pilot's panel, P1 and P3.
Speaker
There is one speaker for the system. It is located on the overhead P5 panel.
Landing Gear Lever Switch
The landing gear lever switch is located in back of the center P2 panel.
Flap Position Switches
The flap position switches are mounted on the flap control unit which is located on the
upper aft portion of the right main gear wheel well.
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GROUND PROXIMITY WARNING – SYSTEM
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COMPONENT FUNCTIONAL DESCRIPTION
CONTROL PANEL
1. Purpose
The ground proximity warning control panel provides the flight crew with visual indication
of GPW operation, self-test capability and flap-gear inhibit capability.
2. Location
The control panel is located on the first officer's P3 panel.
3. Features
The amber INOP light is illuminated when a computer or input signal malfunction is detected,
or a GPWS self-test is being performed.
The FLAP/GEAR INHIBIT switch is a two-position toggle switch, guarded and safety-wired in
the NORMAL position. When it is placed in the INHIBIT position, the TOO LOW-FLAPS and TOO
LOW-GEAR warnings are inhibited.
The SYS TEST switch is used to initiate a GPWS self-test. Self-test can be conducted on the
ground or in-flight.
CONTROL PANEL
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
WARNING LIGHTS
1. Purpose
The warning lights provide visual indication of ground proximity warning.
2. Location
Two warning lights are provided on each pilot's control panel.
3. Features
The red PULL UP warning lights illuminate when a mode 1, 2, 3 or 4 flight path is detected.
The amber BELOW GS warning lights illuminate when glide slope deviation becomes excessive.
Press the BELOW GS lens cap to inhibit its warning. Remove the lens cover to gain access to
the lamps.
WARNING LIGHTS
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COMPONENT FUNCTIONAL DESCRIPTION
COMPUTER
1. Purpose
The computer monitors the input parameters to determine if an unsafe flight condition exists
and provides an appropriate output to the aural and visual warning devices.
2. Location
The computer is located on electronic equipment shelf E2-4.
#. Physical Description
The computer is housed in a rack-mounted, 1/4 ATR short enclosure weighing 8 pounds, and
secured by a single hold-down device. It requires 115 volt ac electrical power and all
electrical interface is provided by two rear-mounted connectors.
4. Operation
The computer processes radio altitude, barometric altitude rate, glide slope deviation, mach
number, landing gear and flap logic which are provided by other airplane systems. If the
airplane enters an unsafe flight path relative to the ground, the computer issues a warning.
5. Self-Test
There is no self-test feature available at the computer.
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COMPUTER
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NOTES
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OPERATION
WARNING MODES
Operation Sequence
The ground proximity warning system provides the pilots with an aural and visual warning of
potentially dangerous flight paths. The computer has six modes of operation.
Description
When the computer issues a mode 1, 2, 3 or 4 warning, the PULL UP warning light illuminates.
When the computer issues a mode 5 warning, the BELOW G/S warning light illuminates.
Aural warnings are provided through the GPWS speaker for all modes. The computer is capable
of providing nine different messages. Each aural warning message has a priority level
assigned to it.
Each of the six modes has a warning envelope and aural messages associated with it.
MODE
1
2
3
4
5
6
PRIORITY
1
2
3
4
5
6
7
8
9
FLIGHT CONDITION
EXCESSIVE DESCENT RATE
EXCESSIVE CLOSURE RATE
DESCENT AFTER TAKEOFF
PROXIMITY TO TERRAIN
DESCENT BELOW GLIDESLOPE
DESCENT BELOW MINIMUMS
MESSAGE
MODE
WHOOP WHOOP - PULL UP
TERRAIN
TOO LOW – TERRAIN
TOO LOW – GEAR
TOO LOW – FLAPS
MINIMUMS
SINK RATE
DON'T SINK
GLIDESLOPE
1 & 2
2
4
4A
4B
6
1
3
5
MODES OF OPERATION
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MAINTENANCE TRAINING MANUAL
OPERATION
GPWS - MODE 1
1. Operation
The purpose of mode 1 is to alert the flight crew to an excessive descent rate with respect
to terrain clearance. This mode is independent of landing gear and flap positions.
2. Normal Sequence
Mode 1 indications occur below 2450 feet radio altitude down to 30 feet when the descent
rate exceeds a threshold value as indicated on the graph. The Mode 1 envelope is divided
into two areas: The initial penetration area (alerting areas, or "SINK RATE" area) and the
inner warning area, "WHOOP WHOOP PULL UP" area. The specific initial penetration area and
the inner warning area boundaries are as shown on the graph.
Initial penetration of the Mode 1 envelope is annunciated by the illumination of the red
PULL UP light and the repeated aural message "SINK RATE....". Penetration of the inner
warning area is annunciated by the red PULL UP light and the repeated aural message "WHOOP
WHOOP PULL UP...."
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GPWS MODE 1
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OPERATION
GPWS - MODE 2A
1. Operation
The purpose of mode 2 is to alert the flight crew to an excessive closure rate with respect
to rising terrain.
2. Normal Sequence
A.
Mode 2 Definition
This mode consists of two submodes: If the flaps are 15 units or less, mode 2A is
annunciated; if the flaps are down greater than 15 units, mode 2B is annunciated. Mode
2B is described later.
B.
Mode 2A Envelope
If the airspeed is less than 220 knots, the upper boundary is 1650 feet radio altitude
and the lower boundary 30 feet. If the airspeed exceeds 310 knots, the upper boundary
is 2450 feet radio altitude and the lower boundary 30 feet. Between 220 knots and 310
knots, the upper boundary varies linearly with respect to computed airspeed (airspeed
expansion function).
Penetration of the mode 2A envelope can be either on the slope or from the top. The
envelope is divided into two areas: the initial penetration area (alerting or
"TERRAIN-TERRAIN..." area), and the inner warning area ("WHOOP WHOOP PULL UP" area).
The inner warning area is entered 1.6 seconds after entering the initial penetration
area.
C.
Mode 2A Indications
Initial penetration area indications consist of the illumination of the red PULL UP
light, and the aural message "TERRAIN". The indications of the inner warning area are
the illumination of the red PULL UP light, and the repeated aural message "WHOOP-WHOOP
PULL-UP"...".
D.
Altitude Gain Function
Upon leaving the inner warning area, due to either terrain drop-off or a pull-up
maneuver, the altitude gain function is activated. During this function, the
indications change to the alerting message - the red PULL UP light and the repeated
aural message "TERRAIN...". The indications continue until the airplane has gained 300
feet of barometric altitude or when the landing gear is lowered.
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GPWS - MODE 2A
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OPERATION
GPWS - MODE 2B
1. Operation
The purpose of mode 2B is to alert the flight crew to an excessive closure rate with respect
to rising terrain with the flaps down greater than 15 units, or the glideslope deviation is
within plus or minus two dots.
2. Normal Sequence
A.
Mode 2B Envelope
The Mode 2B indications occur below 789 feet radio altitude and down to 200-600 feet
depending upon the barometric rate of descent, when the closure rate exceeds threshold
values as shown on the graph.
B.
Mode 2B Indications
(1) Gear Down
When the gear is down, mode 2B indications consist of the repeated alert
"TERRAIN..." and the illumination of the red PULL UP light.
(2) Gear Up
When the gear is up, mode 2B indications consists of the repeated warning "WHOOP
WHOOP PULL UP …” and the illumination of the red PULL UP light.
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GPWS - MODE 2B
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NOTES
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MAINTENANCE TRAINING MANUAL
OPERATION
GPWS - MODE 3
1. Operation Sequence
Ground proximity Mode 3 is designed to alert the flight crew to an excessive loss of
altitude after take-off or during a missed approach.
2. Normal Sequence
A.
Mode 3 Types
Mode 3 provides two types of warnings; one based on barometric altitude and radio
altitude, and the other based on minimum radio altitude clearance during takeoff until
Mode 4 is activated. The visual annunciation of both types is the illumination of the
red PULL UP light.
B.
Alerting
Type-1 Descent after Takeoff ("Don't Sink")
This warning is provided whenever the airplane descends a predetermined amount of
barometric altitude before reaching 667 feet radio altitude. An expanded upper limit
of 1333 feet radio altitude is used at higher airspeeds to prevent premature switching
from Mode 3 to a possible Mode 4 warning on climbout. By delaying the switching from
Mode 3 to Mode 4 to 1333 feet radio altitude at these higher airspeeds, a Mode 4
nuissance warning is avoided.
The allowable barometric descent varies as a function of radio altitude and time.
Type-2 Minimum Terrain Clearance ("TOO LOW TERRAIN")
This warning is based on a minimum terrain clearance, or floor, that increases with
radio altitude during takeoff. A value equal to 75 percent of the current radio
altitude is accumulated in a filter that is only allowed to increase in value. If the
radio altitude should later decrease, the filter will store its maximum attained value.
Further decreases in radio altitude below the stored filter value with gear or flaps up
will result in the warning "TOO LOW TERRAIN".
This warning is provided to prevent a collision, during takeoff climb, into terrain
that produces insufficient closure rate for a Mode 2 warning.
The filter is activated at 150 feet radio altitude on takeoff, and continues until Mode
3 switches to Mode 4, or until the radio altitude decreases below 30 feet.
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OPERATION
C.
Mode 3 Alert Termination
A type -1 warning "DON'T SINK", is terminated when a positive rate of climb is
established. A TYPE-2 warning, "TOO LOW TERRAIN" is terminated when radio altitude is
greater than the stored filter value (floor warning altitude).
D.
Mode 3 Armed
(1) Take Off
The flaps logic within the GPWC is forced to a down state below 30 feet radio
altitude to allow setting the mode select state to Mode 3 during final approach or
when the aircraft is parked on the ground with flaps up.
(2) Approach
During landing approach, Mode 3 is armed when the airplane descends through an
altitude of 245 feet in landing configuration (flaps greater than 15 units, and
landing gear down).
E.
Mode 3 to Mode 4 Switchover
The minimum terrain clearance filter is used to control the switching from Mode 3 to
Mode 4. After takeoff, Mode 4 will normally be enabled when the filter value exceeds
500 feet. This will occur at or above 667 feet radio altimeter, depending on the time
allowed to charge the filter. This applies to airspeeds below 190 kts. For computed
airspeeds greter than 250 kts the filter value must exceed 1000 feet for a Mode 3 to
Mode 4 transition.
Mode 3 selection within the GPWC must be retained during power loss to prevent
inadvertent reselection of the wrong mode (Mode 3 or Mode 4). Mode selection logic is
arranged so that a GPWC in either state can be installed in an aircraft on the ground
and the selection of Mode 3 will be made.
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GPWS - MODE 3
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NOTES
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OPERATION
GPWS - MODE 4
1. Operation
The purpose of mode 4 is to alert the flight crew to insufficient terrain clearance when not
in landing configuration.
2. Normal Sequence
A.
Mode 4 Definition
Mode 4 usually applies during the landing phase of flight. It is armed after takeoff
when the mode 3 critical alerting upper boundary exceeds the mode 4A upper boundary.
It is annunciated in the event of insufficient terrain clearance when the airplane is
not in the proper landing configuration. Mode 4 consists of two submodes. When the
landing gear is up, mode 4A is annunciated. When the landing gear is down, but the
flaps are less than or equal to 15 units, mode 4B is annunciated.
With the flaps less than or equal to 15 units and the gear down, the GPWC switches
automatically to mode 3 when the radio altitude drops below 30 feet, so that mode 3
would be armed at takeoff.
B.
Mode 4A - Gear Up
(1) Envelope
At airspeeds less than 190 knots, the threshold radio altitude is 500 feet. Above
190 knots, the mode 4A threshold radio altitude increases linearly, and above
250 knots, it is 1000 feet.
(2) Indications
In the high-airspeed area, the aural message is the repeated "TOO LOW - TERRAIN".
In the low-airspeed area, it is repeated "TOO LOW GEAR". The visual indication in
both areas is the illumination of the red PULL UP light
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OPERATION
C.
Mode 4B - Gear Down, Flaps ≤15 Units
(1) Envelope
At airspeeds less than 159 knots, the threshold altitude is 245 feet radio
altitude. Above 159 knots, the threshold radio altitude increases linearly.
Above 250 knots, the threshold radio altitude is 1000 feet, as in the case of mode
4A. The lower boundary of the mode 4B envelope is 30 feet radio altitude at all
airspeeds.
(2) Indications
In the high-airspeed area, the aural message is the repeated "TOO LOW-TERRAIN".
In the low-airspeed area it is the repeated "TOO LOW-FLAPS". The visual
indication in both areas is the illumination of the red PULL UP light.
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GPWS - MODE 4
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OPERATION
GPWS - MODE 5
1. Operation
The purpose of mode 5 is to alert the flight crew to an excessive deviation below the
glidepath when making a front-course approach. Mode 5 is armed when the airplane enters
the mode 5 envelope with the gear down.
2. Normal Sequence
A.
Mode 5 Envelope
The mode 5 envelope is divided into two alerting areas: the low-level initial
penetration alerting area, and the normal-level inner alerting area.
The low-level alerting area indications occur below 1000 feet of radio altitude and
down to 30 feet, when the glide slope deviation exceeds 1.3 dots (0.46 degrees). The
specific area boundaries are as shown on the graphic.
The normal-level alerting area indications occur below 300 feet radio altitude
and down to 30 feet, when the glide slope deviation exceeds 2.0 dots (0.7 degrees).
The specific area boundaries are as shown on the graphic.
The additional boundary variations below 150 feet are provided to prevent nuisance
messages by allowing for normal beam variations near the lower envelope threshold.
B.
Mode 5 Indications
Mode 5 is annunciated by the illumination of the amber BELOW G/S light/switch and the
repeated aural message "GLIDE SLOPE". The sound level in the normal-level alerting area
is the same as in modes 1 thru 4, and it is 6 db lower in the low-level alerting area.
The "GLIDE SLOPE" message is repeated more rapidly as the terrain clearance decreases
and/or the glide slope deviation increases.
C.
Cancellation of Mode 5 Indications
Mode 5 aural and visual indications may be inhibited or cancelled by pressing the BELOW
G/S light/switch whenever mode 5 is armed.
If the switch is pressed before the mode 5 indications have started, the visual and
aural indications will be inhibited. If the switch is pressed after the indications have
started, the visual and aural indications will be cancelled. Once cancelled, or
inhibited, the indications cannot be reinstated or rearmed simply by a repeated switch
actuation. Mode 5 is automatically rearmed when the airplane exits from the mode 5
envelope, or if the landing gear is cycled.
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GPWS - MODE 5
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OPERATION
MODE 6 - DESCENT BELOW MINIMUMS
Operation Sequence
When the airplane descends below the MDA (minimum descent altitude) or DH (decision height)
selected on low range radio altimeter No. 1, mode 6 warning is issued.
Description
The low range radio altimeter provides a ground to the computer when the airplane descends
below the MDA or DH. The computer issues a voice alert, MINIMUMS-MINIMUMS, but none of the
ground proximity warnings lights illuminate. This mode is enabled between absolute
altitudes of 1000 feet and 50 feet.
MODE 6 –DESCENT BELOW MINIMUMS
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OPERATION
GROUND PROXIMITY WARNING SYSTEM BLOCK DIAGRAM
Operation Sequence
The ground proximity warning system block diagram shows the ground proximity computer, the
systems that provide inputs and the computer outputs.
Power
The GPWS computer receives 115 volts ac from the captain's load control panel (P18).
System Description
The GPWS computer receives input signals from the LRRA system, the ADC system and the
VHF NAV system. Gear lever and flap position logic is also provided to the computer.
The computer processes these analog and logic signals to detect the airplane's flight
conditions.
The LRRA system provides the GPWS with radio altitude (height). LRRA signals are used
in the control of all modes. Mode 2 requires a radio altitude rate signals (terrain
closure rate) which is derived by the computer, from the radio altitude signal. The
decision height trip is used in the mode 6 warning.
The ADC system provides mach and barometric altitude rate signals to the GPWS computer.
Altitude rate is used to the compute mode 1 conditions, and is used to derive altitude
loss to compute mode 3 conditions. The mach signal is used to set warning thresholds
fo mode 2 and mode 4 computations.
The VHF navigation system provides glide slope deviation signal which is used in the
mode 5 computation.
The flap position switch provides a ground to the GPWS computer, when the flaps are not
in a landing configuration. The landing gear lever switch provides a ground when the
landing gear control lever is in the down position. These switches arm the TOO LOWFLAPS and TOO LOW-GEAR warnings of mode 4. The Flap/Gear Inhibit switch, in the INHIBIT
position, simulates a flap and gear down configuration, inhibiting the flap and gear
warnings.
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OPERATION
System Description (Cont.)
When the SYS
and the INOP
(priority 9),
WHOOP WHOOP –
TEST switch is pressed, the PULL UP, light the BELOW G/S warning lights
failure monitor light illuminate. An aural warning consisting of GLIDE SLOPE
then WHOOP WHOOP - PULL UP is generated. The PULL UP warning lights flash and
PULL UP aural warning is repeated up to 7 times.
When the BELOW G/S warning light/inhibit switch is pressed, with the radio altitude
below 1000 feet, the G/S warning is inhibited. The inhibit is removed when the
airplane climbs above 1000 feet.
The ground proximity failure monitor INOP light illuminates during a system self-test, or
when the GPWS computer, LRRA 1, or ADC 1 fails. The INOP light also illuminates when the
NRRA 1 system is tested.
Aural warnings are applied to the GPW speaker.
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GROUND PROXIMITY WARNING SYSTEM - BLOCK DIAGRAM
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MAINTENANCE PRACTICES
GROUND PROXIMITY WARNING SYSTEM - SELF-TEST
Test Preparations
In order to perform the GPWS self-test, the following systems must have electrical power
applied:
Ground Proximity Warning System
Air Data Computer - 1
Low Range Radio Altimeter - 1
VHF Navigation System
Allow at least 30 seconds to elapse before repeating a self-test.
Functional Test-Airplane on the Ground
Gear control lever must in DOWN position, and flaps not in landing configuration (less than
30 units).
Alert the ground crew, if the flaps are to be moved.
Press and hold the SYS TEST switch on the GPWS control panel.
Both pilots' BELOW G/S lights and the INOP light illuminate, and the aural warning
GLIDE SLOPE! Is emitted one time.
Both pilots' BELOW G/S and PULL UP lights and the INOP light are illuminated, after the
GLIDE SLOPE! warning is emitted, and the aural warning WHOOP! WHOOP! PULL-UP! is
repeated for 5-10 seconds. Then the test is terminated by the ground proximity
computer.
After the first WHOOP! WHOOP! PULL-UP! WARNING, the test has been completed and the
test can be terminated by releasing the STS TEST switch.
Functional Test-Airplane In-Flight
In-flight, the airplane must be more than 1000 feet above the terrain and the landing gear
control lever cannot be in the DOWN position.
The test sequence is the same as it is on the ground except that the test continues as long
as the SYS TEST switch is pressed.
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GROUND PROXIMITY WARNING SYSTEM – SELF TEST
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INTRODUCTION
GROUND PROXIMITY WARNING
1. Purpose
The purpose of the Sundstrand-MARK V ground proximity warning system (GPWS) is to alert the
flight crew to the existence of an unsafe condition due to terrain proximity, and to provide
a warning when windshear conditions are encountered.
2. System Description
A.
GPWS Modes
The various hazardous conditions that can be encountered in flight are divided into the
following modes:
Mode 1 - Excessive descent rate.
Mode 2 - Excessive closure rate with respect to rising terrain.
Mode 3 - Excessive altitude loss during climb-out (in takeoff or during go-around)
when not in landing configuration (landing gear up and/or flaps less
than 25º).
Mode 4 - Insufficient terrain clearance when not in landing configuration (landing
gear up and/or flaps less than 25º).
Mode 5 - Excessive deviation below glide slope when making a front course approach
with the gear down.
Mode 6 - Descent through a selected decision height with the landing gear down.
Windshear mode - Whenever windshear conditions are encountered.
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
GROUND PROXIMITY WARNING SYSTEM
1. General Subsystem Features
The main component of the GPWS is the ground proximity warning computer (GPWC). It
establishes the conditional limits associated with mode annunciation and compares them with the
airplane flight status determined on the basis of received flight parameter inputs. It then generates
the appropriate signals for aural and visual mode annunciation and transmits them to the
respective annunciators.
Sometimes, in emergency situations, when the flight crew is well aware of the prevalent
conditions, it may be desirable to override a flap-up position in order to avoid unnecessary
messages. This may be accomplished by switching on the FLAP INHIBIT switch.
Flight compartment self-tests can be initiated by actuating a self-test switch.
Malfunctions are recorded 1n a ncn-volatile fault memory by flight segments. They are displayed
in a GPWC front panel window when a PRESENT STATUS or a FLIGHT HISTORY readout is
initiated.
2. System Interfaces
A.
Flight Paraneter Inputs
Computed airspeed, barometric altitude (REF.29.92 in.), baro corrected altitude No. 1,
barometric altitude rate, inertial vertical velocity, radio altitude, magnetic track, localizer
deviation, glide slope deviation, latitude, and longitude are the parameters received from
interfacing systems and used to compute the airplane's status with respect to ground
proximity warning system modes.
B.
Outputs
Output signals for the activation of visual indicators consist of discretes transmitted to
illuminate the red PULL UP lights, the red WINDSHEAR lights, and the amber BELOW
G/S lights. The same discretes are transmitted to the flight recorder system.
Audio signals are transmitted to the audio integrating system. In case of failure for either
GPWS or WINDSHEAR, discretes are sent to INOP light.
3. General Operation
A.
Power-On
The GPWC becomes operational when power is applied. The operation is automatic.
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
B.
Flap and Landing Gear Position
Information on flap and landing gear position is
modes. Flap-down and gear-down positions can be
INHIBIT switch. When the switch is actuated, it
and gear-up position information and to simulate
gear-down condition, respectively.
used to arm or to inhibit certain
simulated by actuating the FLAP/GEAR
provides signals to override flaps-up
a flaps-down 25ø or greater, or a
The input signal for landing gear position is "LANDING GEAR LEVER" position.
C.
Flight Compartment Self-Test
The input from the ground proximity SYS TEST switch initiates a flight compartment
confidence or a full vocabulary self-test. The confidence self-test can be initiated
in air or on the ground. The full vocabulary self-test can be initiated only on the
ground.
D.
Glide Slope Mode Inhibit
A discrete from the BELOW G/S light/switch, when actuated, cancels or inhibits the
aural and visual mode 5 (below glide slope) indications. Mode 5 also is automatically
inhibited during a backcourse landing approach.
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GROUND PROXIMITY WARNING SYSTEM
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COMPONENT FUNCTIONAL DESCRIPTION
GROUND PROXIMITY WARNING COMPUTER
1. Purpose
The ground proximity warning computer (GPWC) establishes the limits for the ground proximity
warning system modes and compares the airplane flight and terrain clearance status against
the established mode limits. If the airplane is found to have entered a ground proximity
warning system mode, the computer issues appropriate warning or alerting signals. The
computer stores failure data in a non-volatile memory for display in a front panel window.
2. Location
The GPWC is located in the main equipment center.
3. Physical Description
The BITE display is used for present BITE status (PRESENT STATUS) and past BITE flight
fault history (FLIGHT HISTORY) readouts. It consists of eight LED characters.
The BITE displays is initiated by activating the STATUS/HISTORY switch. It is a threeposition toggle switch, with a spring-loaded center position. To initiate the BITE readout,
the switch must momentarily be placed to either PRESENT STATUS or to FLIGHT HISTORY.
NOTE: If it is desired to prematurely terminate a long flight history readout, the front
panel STATUS/HISTORY switch must be positioned to PRESENT STATUS and held until
message CANCEL appears. This terminates the flight history display with the
message END TEST.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
A.
Specifications
The GPWC is MCU size 2, weighs 5.3 pounds (2.4 Kg), and requires 115 volts, 400 Hz,
single-phase power. It uses forced-air cooling.
B.
Electrostatic Sensitivity
The GPWC is an electrostatic sensitive device and must be handled according to
maintenance manual practice 20-41-01.
4. Control
Five different aural message configurations are available. Selection of a desired
configuration is accomplished by means of four program pins (C6, D6, A7, and D5) in the
mating connector. If none of the pins is grounded, the configuration is "basic". By
grounding the appropriate pin, "alternative audio alert select" (AAAS) #1, #2, #3, or #4 can
be selected. No more than one pin can be grounded at any one time. If more than one pin is
grounded, the configuration becomes "basic".
5. BITE
A.
Purpose
The purpose of BITE is to perform an internal check of computer functions, to
record past faults that occurred during the last ten flights, and to annunciate system
status information.
B.
BITE Tests
(1) Continuous Tests
Continuous tests are completely performed during each program loop. Functions
checked are CPU operation and data input integrity for shorts to ground or open
circuits. Also the air data, IRS, ILS, and radio altimeter systems and the
internal power supply are monitored for valid and/or non-computed data.
(2) Periodic Tests
Test requiring excessive processing time are subdivided into smal segments. Tests
on the individual segments are performed sequentially, one segment during each
program loop. Periodic tests include check on processor instruction sets, program
memory contents, RAM addressing and storage functions, voice memory addressing and
contents, parity check of received data and ability to read the data.
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
(3) Event-Initiated Tests
These test are performed during or after a specific events has occurred. They
include: resetting the program a fraction of a second prior to a power supply
failure; checksumming the data stored in the non-volatile fault memory at powerup; checksumming the data written after entering data; sampling and storing
program pin status at power-up; restating the CPU at a known location in the
program after loss of CPU.
C.
Fault Recording
Faults area recorded in a non-volatile fault memory by flight segments. The beginning
and the of each flight segment are identified using radio altitude, airspeed, and
mode 3-to-4 transitions. Up to 24 faults may be recorded during each flight segment.
GROUND PROXIMITY WARNING COMPUTER
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
WARNING LIGHTS
1. Purpose
The warning lights provide visual indication of ground proximity warning.
2. Location
Three warning lights are provided on each pilot's control panel.
3. Features
The red WIND SHEAR lights illuminate whenever windshear conditions are detected. The red
PULL UP warning lights illuminate when a mode 1, 2, 3, or 4 flight path is detected. The amber
BELOW GS warning lights illuminate when glide slope deviation becomes excessive. Press the
BELOW GS lens cap to inhibit its warning. Remove the lens cover to gain access to the
lamps.
WARNING LIGHTS
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NOTES
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
GPWS FLIGHT COMPARTMENT SELF-TEST
1. Fault Isolation
The GPWS flight compartment self-tests provide a means to check out the system and to
isolate any faults present to the LRU.
If the flight compartment self-test indicates the presence of malfunction, the GPWC selftest readout (initiated from GPWC front panel and discussed earlier in the course) will
point to the faulty LRU.
2. Test
A.
Flight Compartment Self-Test Description
The purpose of the flight compartment self-tests is to check the operation of the
system. Two types of flight compartment self-tests are available: the confidence selftest and the full vocabulary self-test. Both flight compartment self-tests are initiated
by actuating the ground proximity SYS TEST switch on the P3-7 panel.
B.
Flight Compartment Confidence Self-Tests
The flight compartment confidence self-test can be performed either in flight, or on
the ground.
(1) Airborne Flight Compartment Confidence Self-Test
(a) Test Conditions
For the airborne flight compartment confidence self-test to be initiated, the
landing gear must be retracted, and the radio altitude must exceed 1000 feet.
(b) Test Initiation
The test is initiated by actuating the ground proximity test switch for less
than 5 seconds. The test is initiated by the moment the switch is actuated.
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
(c) Test-Pass Annunciation
Upon test initiation, the aural message GLIDE SLOPE is voiced and the amber
alerting light BELOW G/S light and the INOP light are illuminated. This is
followed by the aural message WHOOP WHOOP PULL UP and the red PULL UP light
and INOP light illuminated.
(d) Test-Fail
If a fault is present, only the INOP light will illuminate when the test is
initiated.
(2) Flight Compartment Confidence Self-Test on the Ground
(a) Test Conditions
For the flight compartment confidence self-test to be initiated on the
ground, the landing gear must be down, and the radio altitude must be less
than 30 feet.
(b) Test Initiation
As in the case of the airborne flight compartment confidence self-test, this
test on the ground is initiated by actuating the SYS TEST switch for less than
5 seconds. The test is initiated at the moment the switch is actuated.
(c) Test-Pass Annunciation
The sequence of messages and lights annunciated during the flight compartment
confidence self-test on the ground is identical to the one annunciated
during the airborne flight compartment confidence self-test, except that
after test initiation, certain internal computer checks are executed. The
duration of these checks does not exceed 4 seconds.
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
C.
Flight Compartment Full Vocabulary Self-Test
(1) Test Conditions
For the flight compartment full vocabulary self-test to be initiated, the landing
gear must be down and the radio altitude must be less than 30 feet.
The full vocabulary self-test cannot be performed if a condition exists that
inhibits the confidence self-test.
(2) Test Initiation
The test is initiated by actuating the SYS TEST switch for more than 5 seconds.
The test may also be initiated by actuating the test switch during the "WHOOP
WHOOP PULL UP" annunciation of the confidence test.
(3) Test-Pass Annunciation
Upon test initiation, the same internal computer checks are executed as the ones
mentioned in the flight compartment confidence self-test on the ground.
Subsequently, the confidence test aural and visual indications are annunciated,
followed by all the available aural messages voiced in the order shown in the
graphic.
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
D.
Test-Fail Annunciation
(1) Regular Faults
If a fault is present, both flight compartment self-tests are inhibited. The INOP
light illuminates whenever SYS TEST is pressed.
(2) Latching Faults
On the ground, the flight compartment self-tests also are inhibited if one or more
latching faults occurred during the last flight (as described under GPWC SELF-TEST).
In order to unlock the GPWC, follow the procedure indicated on the graphic.
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MAINTENANCE TRAINING MANUAL
GPWS - FLIGHT COMPARTMENT SELF-TESTS
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GPWC - SELF-TESTS (SHEET 1)
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MAINTENANCE PRACTICES
GPWC SELF-TEST
1. Functional Test
A.
Computer Self-Test Description
The computer self-tests check the status of the computer and provide a front panel
readout of either its present status or of faults that occurred during the last ten
flight segments as recorded in the fault memory. No test results - neither aural
messages, nor visual indications - are annunciated in the flight compartment.
B.
Computer PRESENT STATUS Self-Test
(1) Initiation
The PRESENT STATUS self-test is initiated by momentarily placing the STATUS/HISTORY
switch to PRESENT STATUS. The same internal ground proximity warning computer
checks are executed as during the flight-deck full vocabulary or confidence selftests initiated on the ground.
(2) Test Results
The test results are displayed in the front panel BITE display window. Each
annunciation starts with an eight-character test pattern, and ends with the words
"END TEST".
(a) Normal Display
If no faults have been detected, the display reads "SYSTEM OK". For the
complete readout see GPWS SELF-TEST SHEET 1.
(b) Fault Display
If there are faults, the display annunciates the detected faults; such as,
"RADIO ALTITUDE INVALID", "ILS DATA INACTIVE", "IRS DATA INACTIVE", etc. A
full list of available messages and of possible causes of the faults, is
shown on the two GPWC SELF-TEST graphics.
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MAINTENANCE PRACTICES
2. Operational Checkout
A.
Computer FLIGHT HISTORY Self-Test
(1) Initiation
The computer FLIGHT HISTORY self-test is initiated by momentarily placing the
STATUS/HISTORY switch to FLIGHT HISTORY.
(2) Test Results
(a) Flight Segments
The fault readout includes up to 24 faults that occurred during any of the
last ten flight segments. A flight segment starts when the airplane is
airborne (RA greater than 65 feet and airspeed less than 90 knots). A flight
segment ends when GPWC has transitioned from mode 3 to mode 4, and
subsequently back from mode 4 to mode 3, and the airplane returns to the
ground (airspeed less than 90 knots and RA less than 30 feet). No faults can
now be written into the fault memory.
Flight segment numbering begins with flight 1 as the most recent flight
segment.
Every flight history readout displays the message QNH SELECTED (corrected
barometric altitude to sea level) or QFE SELECTED (airfield elevation) prior
to END TEST.
The flight segment history data is capable of being read on the computer
front panel BITE display at any time without destroying the data.
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
(2) Test Results (Cont)
(b) Normal Display
When no faults were stored in the fault memory, the display reads "PREVIOUS
TEN FLIGHTS OK".
(c) Fault Display
If there are faults recorded in the fault memory, the BITE display scrolls
the recorded faults from right to left; such as, "GPWS FAILED", "RADIO
ALTITUDE INVALID", or "AIR DATA INACTIVE", and the flight numbers during
which these faults occurred. A full list of possible messages is shown on
the GPWC SELF-TEST graphics.
(d) FLIGHT HISTORY Self-Test Termination
If it is desired to prematurely terminate a long flight history readout, the
front panel STATUS/HISTORY switch must be positioned to PRESENT STATUS and
held until the CANCEL message appears on the BITE display. This terminates
the flight history display with the message END TEST.
(3) Computer Failures
Computer failures will be indicated by "GPWC FAILED" or "FLT HIST INVALID" and
will require removal from the airplane for service. All other failure messages
indicate an incorrect input condition and should be corrected without removing the
GPWC.
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MAINTENANCE TRAINING MANUAL
MAINTENANCE PRACTICES
(4) Latching Faults
A number of inflight faults cause the GPWC to lock itself up at landing (Rad. Alt.
Less than 50 feet). Consequently, on the ground the INOP light illuminates and
the flight compartment self-tests are inhibited, even if the faults were
intermittent in flight and are no longer present on the ground. The purpose of
this feature is to cause the maintenance personnel to call up a flight history
readout and to note what faults occurred during the last flight. Subsequently, the
GPWC is then unlocked, the INOP light extinguishes, and the flight compartment
self-test are enable after either the flight history self-test has been performed
or the airplane has taken off (Rad. Alt. Greater than 50 feet).
Some of the latching faults cause the INOP light to illuminate and the flight
compartment self-test to be inhibited in flight, and others do not. The two selftest graphics list both kinds.
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3. MESSAGES CAUSED BY INACTIVE DATA
MESSAGE
RADIO ALTIMETER DATA INACTIVE
AIR DATA INACTIVE
ILS DATA INACTIVE
GLIDSCOPE DATA INACTIVE
IRS DATA INACTIVE
FMC DATA INACTIVE
DWS NO. 1 INACTIVE
DWS NO. 2 INACTIVE
SOURCE OF FAILED
PARAMETER
LEFT RADIO ALTIMETER
LEFT AIR DATA COMPUTER
LEFT DIGITAL-TO-ANALOG ADAPTER
LEFT VHF NAV RECEIVER
LEFT INERTIAL REFERENCE UNIT
FLIGHT MANAGEMENT COMPUTER
LEFT STALL WARNING COMPUTER
RIGHT STALL WARNING COMPUTER
DAA DATA INACTIVE
LEFT DIGITAL-TO-ANALOG ADAPTER
MCP DATA INACTIVE
MODE CONTROL PANEL NO.1
NOTE: 1. CAUSES OF INACTIVE DATA MESSAGES: ABSENCE OF INCOMING DATA,
OPEN OR SHORT CIRCUIT ON DATA BUS, SOURCE OFF OR POWER LOSS
2. INOP LIGHT ILLUMINATED AND CONFIDENCE TEST
INHIBIT WHEN CONDITION PRESENT
56
PRESENT
STATUS
MESSAGE
FLIGHT
HISTORY
MESSAGE
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
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MAINTENANCE TRAINING MANUAL
EFIS Differences
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NOTES
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COMPONENT FUNCTIONAL DESCRIPTION
WARNING LIGHTS
Purpose
The warning lights provide visual indication of ground proximity warning. The windshear
warning message provides visual indication of a windshear condition.
Location
Two warning lights are provided on each pilot's panel (P1 and P3). The windshear warning is
displayed on each pilot's EADI.
Features
The red PULL UP warning lights illuminate when a mode 1, 2, 3, or 4 flight path is detected.
The amber BELOW G/S warning lights illuminate when glide slope deviation becomes excessive.
Pressing the BELOW G/S lens cap will inhibit this warning.
The red WINDSHEAR warning message or the yellow WINDSHEAR caution message is displayed on the
bottom of the EADI.
WARNING LIGHTS
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MAINTENANCE TRAINING MANUAL
OPERATION
WINDSHEAR
System Description
Windshear is a phenomenon which can happen anywhere within the atmosphere. It can have both
horizontal and vertical components, and involves large volumes of air which move rapidly in
different (usually opposite) directions.
One type of windshear which is most threatening to pilots is the microburst, which produces a
column of downward-moving air. Microbursts are most threatening below 500 feet, where
pilots have little time and airspace for recovery. This graphic shows two microburst
situations, one in takeoff and the other on approach, with the appropriate explanations.
Windshear Warning
A windshear warning can occur on takeoff or approach. A windshear warning has priority over
any other GPWS warning.
Windshear Warning Indications
When the ground proximity warning system determines that windshear conditions exist, the aural
warning "WINDSHEAR ..." is repeated and the red windshear warning is displayed on both
EADI's.
Windshear Caution
A windshear caution can occur on approach only and only for headwind/updraft conditions.
Windshear warning, modes 1, 2, 3, 4, and 5 will have priority over a windshear caution.
Windshear Caution Indications
For a windshear caution condition a yellow "WINDSHEAR" will appear on the EADI. No aural
warning is associated with the windshear caution.
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WINDSHEAR
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OPERATION
GPWS - RADIO ALTITUDE AURAL CALLOUTS
Operation
The purpose of radio altitude aural callouts is to alert the flight crew when the airplane is
descending through the corresponding radio altitudes or if a bank angle of 35° is exceeded.
Normal Sequence
Envelope
The radio altitude aural callout function is armed when the airplane climbs out through
1000 feet radio altitude. Once the airplane descends through each callout altitude plus
an additional ten feet, the corresponding aural callout will be disabled. The callouts
are rearmed when the airplane again climbs through 1000 feet radio altitude with a
simultaneous computed airspeed greater than 70 knots.
Indications
There are no light annunciations with the radio altitude aural callouts. As the airplane
descends, the following radio altitude callouts are announced. At 100, 50, 30, 20, and
10 feet radio altitude, the aural callouts are "One-Hundred", "Fifty", "Thirty",
"Twenty", and "Ten", respectively. Each callout is started within ten feet after the
airplane descends through the callout altitude. Callouts may be skipped if they are
inhibited by other higher priority messages in progress when the airplane is descending
through the callout altitude.
At any altitude if a bank angle of 35° is exceeded, "bank angle" will be annunciated.
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GPWS - RADIO ALTITUDE AURAL CALLOUTS
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OPERATION
GROUND PROXIMITY WARNING SYSTEM - MODE ENVELOPE MODULATION
Operation
The purpose of envelope modulation is to modify specific warning and alerting envelopes to
prevent nuisance GPWS mode annunciations in localities with marginal ground proximity
terrain conditions in approach or takeoff.
Envelope Modulation Example
The operation of envelope modulation is illustrated on the graphic by means of an example.
The example chosen is a backcourse approach to Reno, Nevada.
Normal Sequence
Method Used in Envelope Modulation
A number of airports have been identified as having approach or departure peculiarities
that are likely to produce nuisance annunciations. The areas have been defined by means
of latitude and longitude data stored in the GPWC memory. When the GPWC senses the
airplane's approach to such an area, a check of a number of input signals - defined as the
"key" for the given situation - is made to ascertain that the conditions present require
the modulation of the envelopes of one or more GPWS modes. If the conditions are
"right", the GPWC concludes that the key "fits" and consequently proceeds with the
predetermined modulation. If the key does not fit, no envelope modulation takes place.
The Key
The following parameters may be used as conditions for the key.
General GPWS Parameters
The parameters listed below are used in envelope modulation as well as in
conventional GPWS operation:
Radio altitude
Magnetic track
Glide slope deviation
Selected runway heading
Internally generated time
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MAINTENANCE TRAINING MANUAL
OPERATION
Unique GPWS Parameters
The parameters listed below are used in envelope modulation only; they are not
used in conventional GPWS operation.
Baro corrected altitude, referenced to sea level or to runway elevation.
Localizer deviation
Latitude
Longitude
Snapshot
To determine whether or not to modulate the envelope, a "snapshot" check is made.
This check consists in determining the elevation of a "snapshot area" by
subtracting its radio altitude from its corrected baro altitude and comparing the
obtained result with its elevation stored in memory. If both values do not match,
no envelope modulation takes place. The snapshot area is defined by latitude and
longitude and is situated a short distance before the envelope modulation area.
Types of Envelope Modulation
Mode 1
Both warning and alert envelopes are shifted to the right. This allows greater
barometric descent rates before mode 1 annunciations are generated.
Modes 2a and 2b
The upper envelope boundaries are lowered. This allows the airplane to fly closer
to the terrain without generating mode 2 warning and alert annunciations.
Mode 4
The upper boundary of the high-airspeed envelope is lowered to provide a lower
allowable minimum terrain clearance at higher airspeeds.
Mode 5
The upper boundary of the mode 5 envelope is raised and the gear-down requirement
is removed. This allows mode 5 alerts at higher radio altitudes.
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OPERATION
Mode 6
The upper boundary of mode 6 annunciations is raised to the mode 5 boundary, and the
gear-down requirement is removed.
Envelope Modulated Airports
Agana NAS, Guam Is - modes 2A and 4.
Alicante, Spain - mode 2A.
Alice Springs, Australia - mode 2A.
Cairns, Australia - modes 2A and 4.
Canberra, Australia - mode 2A.
Coolangatta, Australia - modes 2A and 4.
Cuenca, Ecuador - mode 2A.
Funchal, Madeira Is - mode 2A.
Geneva, Switzerland - mode 2A.
Hiroshima, Japan - mode 2A.
Hobart, Tasmania - modes 2A and 4.
Hong Kong, B.C.C. - modes 1 and 2A.
Hot Springs, Virginia USA - modes 5 and 6*.
Kagoshima, Japan - modes 5 and 6*.
Launceston, Tasmania - mode 2A.
Leeds/Bradford, U.K. - modes 2A and 2B.
Lisbon, Portugal - mode 2A.
Luxemborg, Luxembourg - modes 2A and 4.
Malaga, Spain - mode 2A.
Melbourne, Australia - mode 2A.
Nice, France - modes 2A and 4.
Nome, Alaska USA - mode 2A.
North Bay, Ontario Canada - mode 2A.
Nurnberg, Germany - mode 2A.
Ontario, California USA - mode 2A.
Paine Field, Washington USA - modes 5 and 6*.
Quito, Ecuador - modes 2A, 2B and 4.
Reno, Nevada USA - modes 2A and 4.
San Diego, California USA - mode 1.
Seoul, Korea - modes 2A and 4.
St. Johns, Newfoundland - mode 2A.
Stephenville, Newfoundland - mode 1.
Taipei (Sungshan), Taiwan - mode 2A.
Tenerife, Canary Islands - modes 4, 5, and 6*.
Unalakleet, Alaska USA - mode 2A.
Vagar, Faroe Islands - modes 2A and 4.
Victoria, B.C. Canada - mode 2A.
Wellington, New Zealand - mode 2A.
Whitehorse, Yukon Territory Canada - mode 2A.
Wrangell, Alaska USA - mode 2A.
Zurich, Switzerland - modes 2A and 4.
*mode 6 is a GPWC option.
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GROUND PROXIMITY WARNING SYSTEM - MODE ENVELOPE MODULATION
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SYSTEM TEST
GPWS FLIGHT COMPARTMENT SELF-TEST
Fault Isolation
The GPWS flight compartment self-tests provide a means for the flight crew to assure themselves
that the ground proximity warning system is ready to be dispatched.
Test
Flight Compartment Self-Test Description
Two types of flight compartment self-tests are available: the confidence self-test and the
full vocabulary self-test. Both flight compartment self-tests are initiated by
actuating the ground proximity SYS TEST switch on the GPWS module in the flight
compartment.
Flight Compartment Confidence Self-Tests
The flight compartment confidence self-test can be performed either in flight, or on
the ground.
Test Conditions
For the airborne flight compartment confidence self-test to be initiated, the
radio altitude must be greater than 1000 feet. The test may be initiated any time
on the ground.
Test Initiation
The test is initiated by momentarily actuating the ground proximity test switch.
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MAINTENANCE TRAINING MANUAL
SYSTEM TEST
Test-Pass Annunciation
Upon test initiation, the aural message GLIDE SLOPE is voiced and the amber
alerting light BELOW G/S light and the INOP light are illuminated. This is
followed by the aural message WHOOP WHOOP PULL UP and the red PULL UP light and
INOP light illuminated. The final step in the confidence test announces WIND SHEAR
and displays the red WINDSHEAR warning on the EADI.
Test-Fail
If a fault is present in the GPWS, usually the INOP light will be illuminated
before the test and will remain illuminated after the test. If a fault is
present, the aural message(s) may or may not be heard during the test.
Maintenance personnel must check the GPWC present status at the ground proximity
warning computer in the electronic equipment compartment for any abnormal
indications during the test.
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MAINTENANCE TRAINING MANUAL
SYSTEM TEST
Flight Compartment Full Vocabulary Self-Test
Test Conditions
For the flight compartment full vocabulary self-test to be initiated, the airplane
must be on the ground.
Test Initiation
The test is initiated by actuating and holding the SYS TEST switch until the "SINK
RATE" aural begins.
Test-Pass Annunciation
Upon test initiation the confidence test is performed, followed by all the
available aural messages voiced in the order shown in the graphic.
Flight Compartment Latching Faults
Some inflight faults are "latched" in the ground proximity warning computer fault
memory. During this condition, the INOP light will remain illuminated as long as the
airplane is on the ground. To remove the latching fault condition, follow the steps on
the graphic.
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GROUND PROXIMITY WARNING SYSTEM – FLIGHT COMPARTMENT SELF TEST
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PRECAUTIONS
CLEAR SURFACES BEFORE OPERATING FLAPS.
ASSURE HYDRAULIC SYSTEMS POWER IS REMOVED.
INSTALL GROUND LOCKPINS IN LAND GEAR PRIOR TO
OPERATING LANDING GEAR LEVER.
USE ELECTRICAL DRIVE TO MOVE FLAPS.
OBSERVE PROCEDURES AND LIMITATIONS OF THE AIR DATA
SYSTEM.
POWER.
PILOT
REPORTS
ON GROUND
TEST
ASSURE POWER IS APPLIED TO THE GPWS AND ASSOCIATED
SYSTEMS
DETERMINES CONDITIONS UNDER WHICH REPORTED
MALFUNCTION OCCURRED.
CHECKS HEIGHT ABOVE TERRAIN LIMITS FOR
GP WARNING.
CONFIRMS VALIDITY OF EXTERNAL SYSTEM.
ACTIVATES THE AUDIO GENERATOR AND EXTERNAL LIGHT
INDICATIONS.
GROUND PROXIMITY WARNING SYSTEM MAINTENANCE SUMMARY
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Instrument
Comparator
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MAINTENANCE TRAINING MANUAL
INTRODUCTION
INSTRUMENT WARNING SYSTEM
1. Purpose
The instrument warning system provides a warning if significant differences exist between the
captain's and first officer's navigation instruments.
2. System Description
The comparator unit receives input from the captain's and first officer's navigation
instruments. If a significant difference exists a warning light is illuminated on the
annunciator panels.
INSTRUMENT WARNING SYSTEM
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MAINTENANCE TRAINING MANUAL
GENERAL DESCRIPTION
GENERAL COMPONENT LOCATIONS
Circuit Breakers
There are two separate circuit breakers on the P6 load control center for the instrument warning
system. One is for the comparator power supply, the other is for the comparator monitor power
supply.
Comparator Unit
The comparator unit is installed on shelf E3-4 in the main equipment center. Connection with
airplane wiring is made through two electrical connectors at the rear of the comparator. A
test switch and several monitor lights are provided on the comparator unit for system test
purposes.
Instrument Warning Annunciator Panels and Test Switch
Instrument warning annunciator panels are located on each pilot's panel. The annunciator
panels contain the warning lights, which illuminate to give a warning indication when the
output of the two systems disagree.
The test switch, on the captain's panel, permits testing of warning channels.
INSTRUMENT WARNING SYSTEM COMPONENT LOCATION
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
COMPARATOR UNIT
1. Purpose
The comparator unit provides a signal output for visual display if a significant difference
exists between the captain's and first officer's navigation instruments.
2. Physical Description
The comparator unit is a 3/8 - ATR short case rack mounted unit. All connections are made
through a plug on the back of the unit. Seven monitor lights are located on the front panel.
The PWR MON light is red and all the others are amber. The amber lights are illuminated by a
comparator warning output.
3. Power
Two 115 volt ac power inputs are required. One input provides power for the instrument
comparators, and the other is used for the power monitor.
4. Operation
The comparator unit is a solid state device consisting of two power supplies and six comparator
monitors. The monitor power supply is used to provide a warning if the comparator power fails.
The comparator monitors heading, bank attitude, pitch attitude, LOC deviation, GS deviation
and radio altitude.
5. Self Test
A PUSH TO TEST switch provides a means of testing the comparator unit. All amber lights
should illuminate when the test switch is pressed, except the power monitor lights, which
are not a part of the self-test. All lights should extinguish when the test switch is
released.
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COMPARATOR UNIT
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
INSTRUMENT WARNING ANNUNCIATOR
1. Purpose
The instrument warning annunciator contains seven amber lights that illuminate to indicate a
discrepancy sensed by the comparator unit.
2. Operation
The MON PWR light illuminates when the instrument comparator power supply fails. The other
monitor lights illuminate when predetermined tolerance thresholds are exceeded. The
tolerance thresholds are listed in the operation section.
The HDG light illuminates when the captain's and first officer's HSI compass cards are
displaying different headings. The PITCH or ROLL lights illuminate when the captain's and first
officer's ADI attitude displays are different. The HDG, PITCH and ROLL annunciators may all
come on simultaneously if excitation voltage from the comparator unit is lost. In such a
case, the Heading, Roll and Pitch comparison is unreliable. The GS and LOC lights
illuminate when the glide slope and localizer deviation signals from the captain's and first
officer's VHF NAV receivers are different.
The ALT light illuminates when the altitude signals from the number 1 and 2 low range radio
altimeter receiver/transmitters are different.
If a light illuminates, it can be dimmed by pressing the clear plastic cover over the lights
(PUSH-TO-DIM).
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MAINTENANCE TRAINING MANUAL
INSTRUMENT WARNING ANNUNCIATOR
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NOTES
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MAINTENANCE TRAINING MANUAL
COMPONENT FUNCTIONAL DESCRIPTION
INSTRUMENT COMPARATOR TEST SWITCH
1. Location
The instrument comparator test switch is a push-button type switch located on the Captain's
panel.
2. Operation
When the test switch is pressed, the lights on both annunciators and the comparator unit
illuminate. The power monitor lights do not illuminate. The lights extinguish when the test
switch is released.
INSTRUMENT COMPARATOR TEST SWITCH
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MAINTENANCE TRAINING MANUAL
OPERATION
COMPARATOR UNIT WARNING THRESHOLD CHART
Control Sequence
The chart shows the threshold values at which the warning lights are turned on by the
comparator unit.
Heading Warning Thresholds
When the airplane is wings level, the heading difference threshold is set at 6° and is
increased to 9° when the airplane is banked. After glide scope capture, the threshold
is reduced to 4°.
Pitch and Roll Warning Thresholds
The pitch and roll threshold is set at 4° difference prior to GS capture and 3° after GS
capture.
LOC Warning Threshold
The LOC difference threshold is large when approaching the LOC center line and decreased
when on the center line. The values are given in the chart.
GS Warning Threshold
The GS difference threshold is large when approaching the GS center line and decreased
when on the center line. The values are given in the chart.
Radio Altimeter Difference Warning Threshold
The radio altimeter warnings are activated at GS capture. The warning threshold is
large at higher altitudes and decreases as altitude is lowered. A step reduction is
made at 500 feet. Interpolation is made between 1500 to 600 and 400 to 0 feet.
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MAINTENANCE TRAINING MANUAL
COMPARATOR UNIT WARNING THRESHOLD CHART
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OPERATION
INSTRUMENT WARNING SYSTEM BLOCK DIAGRAM
Operation Sequence
The comparator unit is the central unit in the instrument warning system. The HSIs, ADIs,
VHF NAV units, radio altimeters and the flight instrument accessory unit provide inputs to
the comparator.
Operation
The comparator unit supplies 26 volt ac excitation to differential resolvers in the
captain's HSI and ADI. The captain's HSI sends heading to the F/O HSI. When the
compass cards' positions do not agree, the F O's HSI develops an error signal and through
the comparator, turns both HDG lights on. The pitch and roll from the ADIs operate in a
similar manner. A separate roll input is applied to the comparator to modify the threshold
at which the heading comparator operates. HDG, ROLL and PITCH MONITOR signals allow the
comparator to sense continuity of the excitation circuits in the HSIs and ADIs.
The VHF NAV
is required
and 28 volt
comparison.
receivers supply LOC and GS deviation for comparison. The 28 volt dc ILS signal
form both receivers to enable the LOC deviation comparison. The GS super flag
dc ILS signals are required from both receivers to enable the GS deviation
When a VHF NAV system is transferred, LOC and GS comparisons are disabled.
The radio altimeters are compared, provided one of the flight control computers (FCC) is in
the GS capture mode. The GS capture signal is also used to lower the threshold for heading,
roll attitude and pitch attitude comparison. The comparator warning threshold chart contained
the values of the thresholds.
The flight instrument accessory unit monitors GS capture status from the FCCs and applies the
status to the comparator unit.
Test
A test switch enables testing of the comparator unit. All lights except MON PWR are turned on
during a successful test.
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INSTRUMENT WARNING SYSTEM BLOCK DIAGRAM
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MAINTENANCE PRACTICES
INSTRUMENT WARNING SYSTEM MAINTENANCE SUMMARY
Functional Test
Refer to the Maintenance Manual for complete troubleshooting, component removal and
installation procedures.
The instrument warning system has inputs from various other systems and these systems must be
operational for complete testing of the comparator warning monitor. The comparator is first
tested by pulling the comparator power circuit breaker which should illuminate the MON PWR
lights. A check is made of the comparator by pressing the Instrument Comparator test
switch.
All lights should come on except the MON PWR lights. While holding test button, press
captain's and first officer's instrument warning panel face. Check that warning lights dim.
Release push TO DIM SWITCH first and then test botton. All lights should extinguish when the
test switch is released
MAINTENANCE
ITEM
ACTION
RESULTS
POWER
ASSURE THAT ALL CIRCUIT
BREAKERS ARE IN TO ALL
INTERFACING SYSTEMS AND TO
THE COMPARATOR UNIT
ALL FLAGS IN THE HSIs, ADIs AND
RADIO ALTIMETERS ARE OUT OF VIEW.
AMBER MON-PWR LIGHT ON INST WARN
ANNUNCIATOR IS OUT.
POWER MONITOR
TEST
PULL COMP PWR CIRCUIT BREAKER
PUSH IN AFTER TEST.
RED MON-PWR LIGHT ILLUMINATES ON
COMPARATOR UNIT AND AMBER MON-PWR
LIGHT ON ANNUNCIATORS ILLUMINATE
COMPARATOR
UNIT TEST
PRESS INST COMP
TEST SWITCH
ILLUMINATE
COMPARATOR
UNIT
DIMMING
TEST
CONTINUE TO HOLD
INST COMP TEST SWITCH IN. PRESS
COVER
OF ONE ANNUNCIATOR
ALL AMBER MONITOR LIGHTS
ALL AMBER INSTRUMENT
WARNING LIGHTS DIM
INSTRUMENT WARNING SYSTEM MAINTENANCE SUMMARY
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