B767 ATA 34 Student Book

TRAINING MANUALFOR TRAINING PURPOSES ONLY

B767-3S2F CH 34-00 Page - 1 28 Mar 2014 EFF - ALL

NAVIGATION CH 34



ATA 34 NAVIGATION SYSTEM TABLE OF CONTENTS

NAVIGATION SYSTEM TOC: ............................................................... 2

NAVIGATION INTRODUCTION ............................................................ 5

PITOT STATIC COMPONENT LOCATIONS ........................................ 9

AIR DATA INERTIAL REFERENCE SYSTEMS.................................. 17

ADIRU COMPONENT LOCATION...................................................... 21

ADIRU INITIALIZATION ...................................................................... 23

ADIRU (ADC) TEST ............................................................................ 43

ADIRU (IR) TEST ................................................................................ 49

ADIRS DISPLAY AND SWITCHING ................................................... 51

ALTERNATE VMO/MMO SELECT SWITCH ...................................... 59

REDUCED VERTICAL SEPARATION MINIMUMS.............................. 63

ANTENNA LOCATIONS....................................................................... 65

RADIO ALTIMETER INTRODUCTION................................................. 67

RADIO ALTIMETER SELF TEST ......................................................... 72

ALTITUDE ALERT SYSTEM................................................................ 75

INSTRUMENT SYSTEM FLIGHT DISPLAY ........................................ 81

EFIS - INTRODUCTION ....................................................................... 85

EFIS EADI AND EHSI DISPLAYS........................................................ 97

VSI OVERVIEW.................................................................................. 101

VOR - INTRODUCTION ..................................................................... 105

VOR SELF-TEST (MEC COMPARTMENT) ....................................... 115

DME - INTRODUCTION ..................................................................... 117

DISTANCE DISPLAY ......................................................................... 123

GPS - INTRODUCTION ..................................................................... 129

GENERAL DESCRIPTION................................................................. 133

GPS COMPONENTS ......................................................................... 137

OPERATION MODES......................................................................... 141

RADIO DISTANCE MAGNETIC INDICATOR (RDMI) ........................ 147

ILS - INTRODUCTION........................................................................ 151

ILS - DISPLAYS.................................................................................. 163

SELF-TEST SEQUENCE ................................................................... 165

MARKER BEACON - INTRODUCTION.............................................. 167

ATC - INTRODUCTION ...................................................................... 173

TCAS - INTRODUCTION.................................................................... 181

TCAS OPERATIONAL TEST.............................................................. 187

WEATHER RADAR - INTRODUCTION.............................................. 193

WEATHER RADAR DISPLAYS.......................................................... 199

PREDICTIVE WINDSHEAR THEORY ............................................... 201

WEATHER RADAR TEST DISPLAYS................................................ 205

EGPWS - INTRODUCTION................................................................ 209

EGPWS - COMPONENT LOCATIONS .............................................. 217

WINDSHEAR MODE .......................................................................... 227

FMCS - COMMUNICATION, NAVIGATION, SURVEILLANCE.......... 231

FLIGHT MANAGEMENT LOCATIONS............................................... 241

MULTIFUNCTION CONTROL DISPLAY UNIT (MCDU) .................... 245

FLIGHT MANAGEMENT COMPUTER INPUTS................................. 249

MAINTENANCE INDEX PAGE........................................................... 261

NAV DATA CROSSLOAD PAGE ....................................................... 263

FMCS - SELF-TEST ........................................................................... 273

TRAINING MANUAL FOR TRAINING PURPOSES ONLY


STUDENT NOTES



NAVIGATION INTRODUCTION

General

Pitot/Static System - senses dynamic (pitot) and ambient (static) air pressure and supplies these to the standby altitude and airspeed indicators and to the air data computers which compute and display air data parameters.

Air Data Computing System - computes airspeed, altitude, mach number and temperature data and supplies it in digital format to interfacing systems.

Air Data Instrument Systems - displays airplane speed and altitude on both pneumatic and electronic displays.

Altitude Alert System - provides aural and visual alert indications when the airplane approaches or departs from a selected altitude.

Inertial Reference System (IRS) - primary reference source for attitude and navigation displays and autoflight systems. The IRS determines and provides angular rates and accelerations and computes attitude, true and magnetic headings, velocity and present position.

Electronic Flight Instrument System (EFIS) - the primary navigation display system utilizes the electronic attitude display indicator (EADI), the electronic horizontal situation indicator (EHSI), radio distance magnetic indicator (RDMI), and the vertical speed indicator (VSI). The EADI and EHSI are CRT displays driven by one of 3 symbol generators. The EADI provides primary attitude, flight director and autoflight mode annunciation. The EHSI is the primary navigation display.

Standby Magnetic Compass - an independent compass providing a backup indication of the airplane's magnetic heading.

Standby Attitude Reference System - provides a backup indication of pitch, roll, and ILS displays.

ILS Navigation System - determines lateral (localizer) and vertical (glideslope) deviations which are displayed by the EADI and EHSI.

Marker Beacon System - provides visual and aural indications when the airplane flies over various types of marker beacons.

Radio Altimeter System - determines airplane height above the terrain for display and use by other systems.

Weather Radar System - displays areas of precipitation ahead of the airplane on the EHSI and a weather radar indicator.

Ground Proximity Warning System (GPWS) - provides aural and visual warnings on the airplane's approach toward terrain, windshear, or departure below glideslope path, by monitoring ILS, IRS, radio altitude, and ADC data.

VOR System - determines bearing with respect to ground-based VOR stations.

Air Traffic Control (ATC) System - derives air data for transmission with selected code and identification to the ground in response to ATC interrogation.

Distance Measuring System (DME) - determines slant range distance from the airplane to DME ground stations and displays it on the RDMI and EHSI.

Flight Management System (FMS) - provides navigation and performance data to the autoflight systems and the flight instrument systems. The FMS continuously calculates and executes optimum airplane performance paths.



NAVIGATION INTRODUCTION



AIR DATA INSTRUMENTS - INTRODUCTION

General

The pitot-static system senses the dynamic (pitot) and ambient (static)air pressure external to the airplane. It supplies these two pressuresto various systems for determining airplane altitude and motion throughthe air mass.

The system consists of aerodynamic compensated pitot-static probes,static ports, drain fittings, and pneumatic tubes and hoses.

The pitot and/or static pressures are supplied to the standby altimeterand standby airspeed indicator. They are also supplied to thedifferential pressure transducer, passenger signs pressure sensor, RATARM Q switch, air data and elevator feel computers.

AIRPLANES WITH INTEGRATED STANDBY FLIGHT DISPLAY;The pitot and/or static pressures are supplied to the Integrated StandbyFlight Display. They are also supplied to the differential pressuretransducer, RAT ARM Q switch, air data and elevator feel computer.

Electrical power is required only for the pitot-static probe anti-icingheaters.



AIR DATA INSTRUMENTS - INTRODUCTION



PITOT STATIC COMPONENT LOCATIONS

Pitot-Static Probe

Two pitot-static probes are installed on the left lower nose section at station 200, and two pitot-static probes are installed on the opposite location.

Each pitot-static probe provides one dynamic and two ambient pressure inputs to various pitot-static subsystems. Pitot pressure is sensed through a single pitot opening at the tip of the probe. Static pressure is sensed through two sets of independent static ports located on the probe. Each pressure source is connected to its respective system.

Each probe is installed with mounting screws with the probe base having two index pins to ensure proper probe alignment. A gasket is installed between the probe base and the airplane structure to form a pressure seal. The probes are not interchangeable with the probes on the opposite side of the airplane.

Heaters are provided for pitot-static probe anti-icing.



PITOT STATIC COMPONENT LOCATIONS



PITOT-STATIC DRAIN LOCATIONS

Pitot-Static System Drains

The pitot-static system drains are concentrated in four areas:

• Forward equipment center (left and right sides of the nose wheel well) • Main equipment center (on stanchions of the electronic racks, left and right

sides) • Aft cargo compartment (on left side, near the door) • Stabilizer compartment (on the left side, behind pressure bulkhead)

The system drain acts as a sump to remove condensation collected from the pitot-static lines. The sump has a reinforced transparent section of tubing with an orange float. This forms a sight gage to indicate the level of liquid accumulated in the sump.

The lower portion of the drain contains a poppet valve covered by a bayonet cap. To drain the pitot static line, the cap is removed and the valve depressor on the cap is inserted into the poppet valve. Accumulated liquid in the sump is drained by gravity flow as the valve is depressed.

Note: The transparenet tubing is not skydrol resistant. If the elevator feel computer leaks skydrol into the static line, the lines must be flushed, and a drain may need to be replaced.



PITOT-STATIC DRAIN LOCATIONS



STATIC COMPONENT LOCATIONS

Alternate Static Port

The alternate static ports are flush mounted on each side of thelower forward fuselage at body station 465. Anti-icing heaters arenot provided on the ports.

Each of the two ports is an independent sensor of external ambientpressure. The static port is cross connected with the port on theopposite side. It provides an alternate source of ambient pressurefor the air data instruments. At the port, pressure is sensedthrough small holes open to the static line tubing.



STATIC COMPONENT LOCATIONS



PITOT STATIC SYSTEM SCHEMATIC

Inputs

The pitot probes sense total pressure. Openings along the pitot probe and the flush-mounted static ports sense ambient (static) pressure. Static sources are cross-connected to compensate for airplane maneuvers.

Distribution

Tubing connects the pressure sensors to pressure-sensitive devices in indicators and computers. The standby altimeter, standby airspeed indicator, and cabin differential pressure indicator convert the air pressures into visual indications. The differential pressure sensor (used to compute cabin differential pressure) is a piezoelectric pressure transducer located below the flight compartment. Pressures are also used in the air data computers and the elevator feel computer. Control circuits for the ram air turbine (RAT) utilizes pitot/ static inputs from the RAT airspeed switch to prevent premature deployment.

Drains are located at low points in the tubing to allow moisture condensation to be removed from the system. The tubing is mounted so that moisture flows down to these drains for all normal airplane attitudes.

Air Data Modules

Air data modules (ADM) connect to pitot and static component to supply digital information to the ADIRU’s.



PITOT STATIC SYSTEM SCHEMATIC



AIR DATA INERTIAL REFERENCE SYSTEM

General

The air data inertial reference system provides air data outputs to the air datainstruments and inertial reference data to other interfacing systems. The system consists of two total air temperature (TAT) probe, two air data inertial reference computers (ADCs), and two angle of attack (AOA) sensors, inertial mode reference panel (IRMP) and air data modules (ADM). The system also has three external test switches



AIR DATA INERTIAL REFERENCE SYSTEM



ADIRU GENERAL DESCRIPTION

General

The air data inertial reference system provides air data outputs to the air datainstruments and inertial reference data to other interfacing systems. The system consists of two total air temperature (TAT) probe, two air data inertial reference computers (ADCs), and two angle of attack (AOA) sensors, inertial mode reference panel (IRMP) and air data modules (ADM). The system also has three external test switches.



ADIRU GENERAL DESCRIPTION



AIR DATA INERTIAL REFERENCE SYSTEM COMPONENTS

Component Locations

• Air Data Inertial reference units - Located on E1-6 • Inertial reference mode panel - Located on P5 • Vertical speed indicators - Located on captain's or first officer's instrument

panels, P1 and P3 • IRS DC power disconnect relay - Located on P6 • AC circuit breakers - One 115 volt ac circuit breaker for each IRU is located

on P11 • DC circuit breakers - One 28 volt dc circuit breaker for each IRU is located

on P6

ADIRU Purpose

The air data inertial reference unit (ADIRU) contains three laser gyros for sensing airplane angular rate about the pitch, roll, and yaw axes, and three accelerometers for sensing linear acceleration along the airplane longitudinal, lateral, and vertical axis. An internal digital computer uses these signals to calculate airplane present position.

Motion Detection

If motion is detected during the align mode, the alignment will automatically be restarted 30 seconds after the motion has stopped. This restarted alignment will require only 8 minutes, omitting the usual 2 minute initial standby time. Also, the initial present position will be either the most recent pilot entry made during the align mode (either before or after the restart) or the last computed position before a down mode alignment if no pilot entry was made.

Auto-calibrate Function

Any change in pitch, roll or yaw caused by biasing errors results in appropriate corrections, rather than continuing to use initial conditions.

Drift Angle Tolerance

External drift angle limited to 60 degrees before the IRU output is tagged invalid.

Latitude Comparisons

During initialization, entered latitude is compared to the last stored latitude immediately following entry, and again with the computed latitude after ten minutes in align mode. In the event of no initial position entry during the 10 minute align period or a miscompare (requiring another entry) is detected, the IRU will flash the ALIGN Light.

Time to Nav

The IRU has the capability to display on the IRMP the time remaining, in minutes, until completion of alignment.

Barometric Altitude Tolerance

The ADIRU barometric altitude reasonableness limit test tolerates a barometric altitude input of (-) 2000 feet to prevent false VSI flags during extremely high barometric pressure periods.

Post-Shop-Visit Position Compare

No initialization position comparison to the last stored position is made during the first alignment cycle following a shop visit by the ADIRU.

Attitude Mode Select Delay

The ADIRU delays actual entry into attitude mode for 2 seconds after selection in order to preclude accidental selection of the attitude mode by overshooting the NAV position of the mode select switch.



AIR DATA INERTIAL REFERENCE SYSTEM COMPONENTS



ADM COMPONENT LOCATIONS

General

There are 7 Air Data Modules located in the Forward end of the main equipment center, just aft of the forward equipment compartment. There are 4 ADM’s on the left and 3 ADM’s on the right side.



ADM COMPONENT LOCATIONS



ADIRU - INTRODUCTION

Purpose

The Air Data Inertial Reference System is one of the primary sensing systems. It provides the primary and navigational parameters indicated in the associated graphic.

System Components

The system consists of three Air Data Inertial Reference Units (ADIRU’s) which sense: angular rates about the X (roll) axis, Y (pitch) axis, and Z (azimuth) axis using laser gyros; and, linear accelerations along the same three orthogonal axes using accelerometers. The inertial reference mode panel (IRMP) provides system control. The ADIRS system conforms to ARINC 704, as well as ARINC 600 and ARINC 429.

Inputs

Air data provides barometric altitude and altitude rate for altitude and vertical speed damping. It also provides true air speed for wind computations. The control display units of the flight management computer system can be used for system initialization.

System Controls

The inertial reference mode panel (IRMP) provides system mode selection, system monitoring and an alternate method of initialization.

System Outputs

A number of airplane systems use the ARINC 429 outputs of the system, including display indicators of the flight instrument system.



ADIRU INTRODUCTION



AIR DATA DIGITAL INPUT/OUTPUT INTERFACES

Inputs

Air data provides barometric altitude and altitude rate for altitude and vertical speed damping. It also provides true air speed for wind computations. The control display units of the flight management computer system can be used for system initialization.

Pitot and Static inputs are sent to Air Data Modules (ADM) and converted to digital signals for the ADIRU.

Outputs

ADIRU outputs are sent for display to the GGU’s the to the Large Display System (LDS). Outputs are also sent to other various systems for processing.



AIR DATA DIGITAL INPUT/OUTPUT INTERFACES



ADIRS THEORY OF OPERATION

Navigation Alignment

During alignment the air data inertial reference system determines the local vertical, the direction of true north, and the initial latitude.

Gyrocompass Process

Inside the inertial reference unit, the three gyros sense angular rate of the airplane. Since the plane is stationary during alignment, the angular rate is due to earth rotation. The ADIRU computer uses the direction of angular rate to determine the direction of true north.

Initial Latitude and Longitude

During the ten minute alignment period, the ADIRU computer has determined true north by sensing the direction of the earth rotation. The magnitude of the earth rotation vector allows the ADIRU computer to estimate latitude of the initial present position. This calculated latitude is compared with the latitude entered by the operator during initialization. Longitude cannot be determined by the ADIRU during alignment

However, the longitude entered by the operator during initialization is compared with the longitude stored in memory the last time the ADIRU was powered down.

Present Position

During initialization, the latitude and longitude of the starting point are entered into the air data inertial reference unit computer. Present position at all future times is determined by adding the distance traveled onto the coordinates of the initial starting position.

Distance traveled is determined by measuring linear acceleration (from the accelerometers) and integrating the result to obtain velocity and integrating again to obtain distance.

Triple Axis Navigation Computation

As long as the airplane flies in only one direction, one accelerometer is sufficient to determine distance traveled from the starting position. Since the airplane may fly in any direction, three accelerometers, mounted to sense acceleration 90 degrees apart, are required.

The three accelerometers are stationary relative to the airplane frame. To determine how much acceleration is causing horizontal movement on the earth,the outputs to the accelerometers have to be compensated by the ADIRU computer, taking into account the airplane attitude and earth curvature.

The compensated outputs from the accelerometers are vector added to determine the actual direction of travel and the amount of travel horizontally. In general, the accelerometers are not oriented north-south and east-west but, their output signals can be related to a north-east coordinate system and the present position can then be determined in terms of latitude and longitude.



ADIRU THEORY OF OPERATION



INERTIAL REFERENCE MODE PANEL

Mode Select Switches

Each inertial reference unit is controlled by its mode select switch on the IRMP. Each mode select switch has four positions:

OFF - removes power from the ADIRU except for logic circuitry associated with the power-off functions.

ALIGN - the inertial reference unit uses earth rotation and gravity to align its reference to the local vertical, to locate true north, and to estimate latitude. Airplane present position must be entered before alignment is completed.

While ALIGN or NAV is selected, the time-to-NAV (TTN) will be displayed in the upper right display window as long as the DSPL SEL switch is in the HDG position. As soon as NAV mode is attained, this display will blank.

NAV - the inertial reference system performs unaided inertial navigation. NAV position has a detent which requires a pull force when switching from NAV to OFF, ALIGN, or ATT. This prevents inadvertent switching from the NAV mode.

ATT - provides rapid attitude and heading restart after total power shutdown to the ADIRU. ATT may also be selected if a fault prevents navigation computations but the ATT mode is still operational.

Mode and Status Annunciators

There are twelve annunciators on the IRMP, four for each ADIRU:

• ALIGN - illuminates white when the ADIRU is in the alignment mode (approximately 10 minutes if the mode select switch is in NAV). It remains illuminated as long as the mode select switch is in ALIGN. The ALIGN annunciator flashes if the alignment procedure fails.

• ON DC - illuminates amber to indicate that the IRU is operating on battery power because 115 volts ac is not available. When the system is initially turned on, the ON DC annunciator illuminates momentarily because the IRU switches off 115 volts ac to verify that battery power is available. This is a normal result of the power-up sequence test done by the ADIRU.

• DC FAIL - illuminates amber when the battery power source drops below 18 volts dc. ON DC and DC FAIL cannot both be on at the same time.

• FAULT - illuminates amber when IRS failures are detected.

DSPL Select Switch

• TK/GS (Track angle/ground speed) - True track angle from 0 through 359.9 degrees is displayed in digits 3 through 6, with a resolution of 0.1 degree. Ground speed from 0 through 2,000 knots is displayed in digits 10 through 13, with a resolution of 1 knot.

Example: 123.4 degrees 321

• PPOS (Present position) - Latitude from 90 degrees S to 90 degrees N is displayed in digits 1 through 6, and longitude from 180 degrees E to 180 degrees W is displayed in digits 7 through 13. Resolution is 0.1 minute.

Example: N89 degrees 59.9' W179 degrees 59.9'

• WIND (Wind angle/wind speed) - True Wind angle from 0 to 359 degrees is displayed in digits 4 through 6 with a resolution of 1 degree. Wind speed from 0 through 256 knots is displayed in digits 11 through 13 with a resolution of 1 knot.

Example: 321 degrees 50

• HDG (Heading) - True heading from 0 to 359.9 degrees is displayed in digits 3 through 6 with a resolution of 0.1 degree. Digits 7 through 13 are blank.

Example: 123.4 degrees

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 variable zero-to-five volt ac signal provided by the aircraft light dimming control circuits.



INERTIAL REFERENCE MODE PANEL



ADIRU INITIALIZATION USING FMC CDU

General

The inertial reference system can be initialized by entering present position on the control display unit of the flight management computers (FMC CDU) or by using the IRMP. Initializing with the FMC CDU or the IRMP requires the data be entered only once for those ADIRU’s currently in the alignment procedure, as indicated by the ALIGN annunciators.

Initialization Procedure (FMC CDU)

Place the ADIRU’s in ALIGN or NAV modes using the mode select switches on the IRMP. Observe that the ON DC annunciator and then the ALIGN annunciator illuminate for all three ADIRU’s.

• Call up the position initialization page on the FMC CDU by pressing the INIT/REF key on the CDU

• SET IRS POS. (5R) will contain box prompts which will allow present position data entry. Use one of the following ways to enter present

• position. If 5R blank, data cannot be entered.

• Scratch pad entry. Enter a latitude and longitude into the scratch pad by pressing the alphanumeric keys. Line select (press) 5R and the scratch pad contents will transfer to the SET IRS POS

• Enter LAST POS. Line select (press) 1R. LAST POS. latitude and longitude appear in the scratch pad. Line select (press) 5R and the scratch pad contents are transferred to the SET IRS POS

• Enter REF AIRPORT. Use alphanumeric keyboard to select the four character ICAO airport identifier. Line select the scratch pad contents into the REF AIRPORT line (2L). Stored latitude and longitude for the airport will be displayed on line 2R. Line select (press) 2R and the latitude/longitude will transfer to the scratch pad. Line select (press) 5R and the scratch pad contents are transferred to the SET IRS POS

• Enter GATE. Line select into 2L a REF AIRPORT as shown previously. Gate identifiers associated with the REF AIRPORT are the only valid entries. Use alphanumeric keyboard to select the GATE (format is 5 characters maximum). Line select (press) 3L to transfer the scratch pad contents to GATE. Stored latitude and longitude will be displayed on 3R. Line select (press) 3R and the latitude/longitude will transfer to the scratch pad. Line select (press) 5R and the scratch pad contents are transferred to the SET IRS POS

Verify the ADIRU’s have accepted the initialization latitude/ longitude by checking the display window on the IRMP with PPOS selected on the DSPL SEL switch. The POS. REF page also will display acceptance of latitude/longitude by the ADIRU’s. To access POS REF, press the NEXT PAGE key on the CDU while the POS. INIT page is displayed. If an ADIRU does not reflect back the entered coordinates within 5 seconds after they were entered, the FMC CDU will display an alert message in the scratch pad, RE-ENTER IRS POSITION.



ADIRU INITIALIZATION USING FMC CDU



ADIRU NORMAL ALIGNMENT PROCEDURE

General

During the alignment process, the ADIRU determines the local vertical and the direction of true north. The airplane cannot be moved during alignment.

Normal Procedure

Alignment can be achieved by the procedure shown with the mode select switches in ALIGN. Normally alignment takes a minimum of ten minutes at which time the ADIRU’s are ready to be switched into the NAV mode. The operator must insert present position sometime during the alignment process using either the FMC CDU or the IRMP. Problems with the alignment process are indicated by a flashing ALIGN annunciator or steady FAULT annunciator on the IRMP.

The alternate alignment procedure is to move the IRMP mode select switch directly into NAV. The ADIRU automatically advances to the navigate mode at completion of the ten minute alignment if present position has been entered. If a problem occurs during alignment, the fault annunciator illuminates, and if present position has not been entered by the time alignment is complete the ALIGN annunciator flashes.

Time-To-Navigation Mode Display

The time interval, in minutes, for an ADIRU to enter the navigation mode may be displayed as depicted on the graphic.

ADIRU INITIALIZATION - USING IRMP

The IRMP can also initialize the inertial reference system. Present position is entered into all ADIRU’s that are aligning, as indicated by ALIGN annunciators. Initialization must occur before the ADIRU’s will complete the alignment process.

Procedure

The Procedures to align are as follows:

Place mode select switches for the ADIRU’s in ALIGN or NAV. Check that the ON DC annunciators illuminate momentarily and then the ALIGN annunciators.

Place DSPL SEL in PPOS position (to verify IRU has been initialized).

Enter the latitude and longitude of present position with the keyboard.

Either latitude or longitude can be entered first.

For latitude, press N2 or S8 key. The letter N or S will appear on the left digit of the left display and the rest of the display will blank.

Continue to enter latitude. As a key is pressed, the digit appears in the right digit of the left display and remaining digits shift one to the left. Press ENT to enter the latitude into the ADIRU computer.

Longitude is entered in the right display in a similar way, starting with the W4 or E6 key. Press ENT to enter 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 mistake is made before ENT is pressed, the CLR key allows the displays to be cleared.



ADIRU NORMAL ALIGNMENT PROCEDURE

ON N101FE IRMP STARTS AT 6 AND GOES DOWNTOTAL TIME TO ALIGN TIMED AT 4 MIN 39 SEC



IR ADVISORY STATUS AND MAINTENANCE MESSAGES

General

To see ADIRU stored maintenance messages on the IRMP panel:

• Select desired ADIRU • Display Switch to “HDG” • Type 01 • Then push “CLR” to advance to the next code. • If no more codes are stored returns to normal displays



IR ADVISORY STATUS AND MAINTENANCE MESSAGES



INERTIAL REFERENCE SYSTEM FMC MESSAGES

General

FMC messages are provided to assist the operator during IRS alignment. These messages (Table 1) are generated using FMC logic and ADIRU digital discretes (ARINC 429 dataword label 270). Whenever an IRS/FMC message is shown, the EICAS level B message FMC MESSAGE is shown and the amber FMC annunciator (PI-3) is illuminated.



INERTIAL REFERENCE SYSTEMS FMC MESSAGES



ADIRU ATTITUDE MODE

General

In the attitude mode, the ADIRU has only limited capability and few outputs, the most important of which is pitch and roll attitude. This mode is entered by moving the mode select switch to the ATT position.

The attitude mode has only limited use. It could be selected in either of two situations. One situation would be if only attitude information is needed. This could occur if a weather radar check is required on the ground.

Another situation would be if the ADIRU navigation functions fail but the attitude functions remain operational. An example of the second situation would happen if an ADIRU had an AC and DC power interruption in flight.

When the attitude mode is selected, the ADIRU is latched into this mode even if the switch is moved to ALIGN or NAV. To select another mode, OFF must be selected first.

Attitude Outputs

When the ATT position of the mode select switch is selected from OFF or ALIGN or NAV, a thirty second alignment period is required. During this time local vertical is sensed. After the alignment period, pitch and roll attitude, accelerations, and inertial vertical speed are output.

Heading Outputs

If ATT is selected and a magnetic heading output is desired, the heading has to be initialized through the FMC CDU or the IRMP. If magnetic heading is initialized, the ADIRU will use this as initial magnetic heading output and will change magnetic heading output as the platform heading changes.



ADIRU ATTITUDE MODE



AIR DATA TEST AND DISPLAYS

ADIRU and Yaw Damper Test Switches

Both the ADIRU and yaw damper test (from the test panel on P61) causes all three ADIRU’s to go into self-test. The test mode is inhibited in the NAV mode when ground speed is greater than 20 knots, and it is also inhibited in the ATT mode.

If possible, turn off the hydraulics to the yaw damper before pushing the ADIRU or the yaw damper test switch.

Performing the ADIRU or yaw damper test with the hydraulic system pressurized causes movement of the rudder assembly.

WARNING: ALL CONTROL SURFACES ARE HYDRAULICALLY POWERED AND MAY MOVE WHEN ANY HYDRAULIC SYSTEMS ARE PRESSURIZED, OR IF ANY CONTROLS ARE MOVED. ALL PERSONNEL AND STANDS SHALL BE CLEAR OF CONTROL SURFACES AND CONTROL COLUMN WHEN HYDRAULIC SYSTEMS ARE PRESSURIZED.

ADIRU Test Switch

The test mode is inhibited in the NAV mode when ground speed is greater than 20 knots, and it is also inhibited in the ATT mode.

Test Results - IRMP

Initiating a self-test causes all annunciators for that ADIRU to illuminate for two seconds. Also, all segments of the display are illuminated for two seconds (except the most significant character of longitude reads 1). After ten seconds, the ADIRU outputs go to preset test values briefly.

The master dim and test switch may be used to test the IRMP instead of the test switch on each individual ADIRU. If the master dim and test switch is to be

used, all three ADIRU’s should be installed or the master dim and test IND LTS switch should be in the BRT position.

CAUTION: DO NOT OPERATE MASTER DIM AND TEST SWITCH FOR MORE THAN FIVE MINUTES WITH MASTER DIM AND TEST IND LTS SWITCH IN DIM POSITION WHEN ANY OF THE THREE ADIRU’S ARE REMOVED. DAMAGE TO THE IRMP CAN RESULT.



AIR DATA TEST AND DISPLAYS



ADIRU (ADC) TEST (TYPICAL)

Test Initiation

There is no test that can be initiated on the ADIRU.

On the P61 test panel:

The identical test for the left or right air data computer is initiated by moving the spring loaded, center-off, toggle switch up to the "L ADC" or "R ADC" position.

Note: Test capability via the test panel on P61 is inhibited in flight.

Test Results

The test results and their sequence of occurrence are as shown on the graphic.



ADIRU (ADC) TEST (TYPICAL)



ADIRU TEST DISPLAYS

EICAS

Several air data parameters are available on the EICAS displays. TAT is always displayed as a primary display. TAT, CAS, MACH and ALT can be displayed on the lower EICAS display unit by pressing the PERF/APU key on the EICAS maintenance control panel (P61).

Because these parameters are found on the EICAS maintenance pages, they are only available on the ground. These parameters are an excellent method for cross-checking TAT, CAS, MACH, and ALT test values during air-data self tests.

FMC

TAS and SAT are shown on the flight management computer control display unit (FMC-CDU) on PROGRESS page 2/2. Pressing the PROG key will display page 1/2. Pressing the NEXT PAGE key will display page 2/2 on which TAS and SAT are located.



ADIRU TEST DISPLAYS



IR SELF TEST AND DISPLAYS

ADIRU Test Switches

Moving the ADIRU test switches down (from the test panel on P61) causes all three ADIRU’s to go into self-test of the inertial reference system. The test mode is inhibited in the NAV mode when ground speed is greater than 20 knots, and it is also inhibited in the ATT mode.

The test will perform the following:

ADI tilt to 45 degrees roll and 5 degrees pitch up

HSI Magnetic heading will go to 15 degrees

IRMP will go to all indicators on for 2 seconds then off for 2 seconds

Depending on what position is selected on the IRMP (TK/GS, PPOS, WIND, or HDG) indications are shown for each display position.

TK/GS - 00 200PPOS - N 22300 E 22300WIND - 30 100HDG - 100



IR SELF TEST AND DISPLAYS



ADIRS DATA DISPLAY AND SWITCHING

Instrument Source Select Panel

ADIRS source select switches allow the captain (P1-1) and first officer (P3-3) to switch between the normal (on side ADIRU) and alternate (center IRU) source of ADIRU data.

Switch push-button illuminates white when switch is in ALTN position.

Radio Distance Magnetic Indicator

The RDMI's display magnetic heading information supplied by the offside selected IRU.

The HDG failure flag appears on the instrument face if the ADIRU magnetic heading is invalid or NCD.

Vertical Speed Indicator

The VSI displays vertical speed from the on side selected ADIRU.

The OFF failure flag appears if ADIRS VSI data is invalid or NCD.

Electronic Horizontal Situation Indicator

The EHSI displays heading information supplied by the selected on side ADIRU. Track and wind data is dynamic data from the selected FMC, but defaults automatically to the ADIRU.

ADIRS failures cause removal of ADIRS-related Symbology and display of the TRK failure flags.

Electronic Attitude Director Indicator

The EADI displays pitch and roll from the selected ADIRU. Ground speed is dynamic data from the selected FMC that defaults automatically to the IRU.

ADIRS failures cause removal of ADIRS-related Symbology and display of the ATT failure flag.



ADIRS DISPLAY AND SWITCHING



ADIRU ALIGNMENT INDICATIONS

General

This table lists the indications visible on the IRMP during the IRS alignment process. In general, the FAULT annunciator is on steady for faults, and the ALIGN annunciator flashes if operator attention is needed.

Longitude Comparison

The ADIRU compares the longitude entered during initialization with the longitude stored in memory of the last position. If the two differ by more than one degree, the ALIGN annunciator flashes immediately. If the ADIRU was newly installed or the airplane ferried without using that ADIRU, this would be a normal display. Entering the longitude a second time forces the IRU to accept the new longitude.

Latitude Comparison

The ADIRU compares the latitude entered during initialization with the last position latitude stored in memory. If the two differ by more than one degree, the ALIGN annunciator flashes immediately.

The latitude entered during alignment is stored until alignment is completed. After the alignment is completed the ADIRU compares latitude calculated with the latitude entered for initialization. If the two do not agree, the ALIGN annunciator flashes. If the same latitude entered the second time still does not compare with calculated latitude, the FAULT annunciator illuminates and the ALIGN light comes on steady.

If the two entries mentioned above were done with wrong latitude values, a subsequent entry of the correct latitude will be accepted by the ADIRU and the fault light and the ALIGN light will extinguish.



ADIRU ALIGNMENT INDICATIONS



ADIRS INPUTS

System Power

Normal system power is 115 volts ac from circuit breakers on the P11 panel with 28 volts dc from the hot battery bus providing a backup power source. For system startup, ac or dc power must be available.

Switching to 28 volts dc is accomplished automatically by the ADIRU’s when loss of 115 volts ac is sensed. Five minutes after 28 volts dc is supplied from the main battery relay, the backup hot battery bus 28 volts dc is removed from the right IRU by the IRS DC power disconnect relay. The center and left ADIRU’s remain powered from the airplane battery. During autoland the center bus isolation relay K123 inhibits the IRS DC power disconnect relay.

Note: After a five minute time delay due to AC power lost the left and center ADIRU’s will shut down with the right continuing on DC power until the aircraft battery is depleted.

ADIRS Inputs

The inertial reference mode panel provides mode select discretes to the ADIRU’s.

The left and right air data computers provide altitude, altitude rate, and true airspeed. For the left and right ADIRU’s, the ADC is selected by the on side ADC instrument source select switch. The center ADIRU receives a switching discrete from the first officer's IRS source select switch to control which (left or right) ADC input it uses. In the normal, position the left ADC supplies the center ADIRU, in ALTN (alternate) position the right ADC supplies the center IRU.

Data also comes from both left and right flight management computers as initialization inputs.

The L/R YAW DMPR test switch on the P61 panel will cause the left or right yaw damper module to go in to test. When either module is in test it sends an ADIRU test discrete to all three ADIRU’s.

Ground Warning

If ac power is lost and any ADIRU is on, the ground crew call horn will sound to warn personnel that the ADIRU is being powered from the airplane battery.



ADIRS INPUTS



ADIRS OUTPUTS

Output Signals

The high speed ARINC 429 data buses transmit data from each ADIRU related to airplane heading, attitude, inertial velocities, position, acceleration, angular rates, and wind speed and direction. Status discretes route to the inertial reference mode panel and to the EICAS computers for display on the upper EICAS display unit.

Interfacing Systems

This sheet shows the ADIRU that provides data to each interfacing system.

MMR Input for GPS initialization is provided on provisional aircraft.



ADIRS OUTPUTS



ALTERNATE CAPT AND F/O’S CADC SELECT SWITCH

Features

An alternate CAPT / F/O’s SELECT guarded and wire-locked switch is provided to allow selection of Center ADIRU for CADC functions during Deferals.

When a Captains (Left) or F/O’s (Right) Air Data function is inop, ALTN can be selected to allow deferral of the Captains or F/O’s and the Center ADIRU would be used for indication. This switch has to be actuated prior to takeoff since it is located in the main equipment center at the outboard side of the E-1 Rack.



CENTER AIR DATA SYSTEM SELECT SWITCHES (MEC)



ALTERNATE VMO/MMO SELECT SWITCH

Features

An alternate VMO/MMO SELECT guarded and wire-locked switch is provided to accommodate a flight with the landing gear extended.

When a flight with the gear down is anticipated, this switch has to be actuated prior to takeoff since it is located in the main equipment center.

Actuation of this switch modifies the air data inertial reference unit software such that the maximum operating speeds allowed are greatly reduced. The specific values are provided on the graphic 767 MAXIMUM OPERATING SPEED SCHEDULE.



ALTERNATE VMO/MMO SELECT SWITCH

The ALTERNATE VMO/MMO select switch allows the airplane to fly with the landing gear extended.VMO = VELOCITY MAXIMUM OPERATINGMMO = MACH MAXIMUM OPERATING



ADIRS MESSAGES

General

The upper EICAS display unit on the EICAS panel announces the IRS ON DC, IRS DC FAIL, and IRS FAULT messages as a level C message. These messages appear at the same time as the amber annunciator lights on the IRMP.



ADIRS MESSAGES



REDUCED VERTICAL SEPARATION MINIMA

Purpose

Reduced vertical separation minima (RVSM) permits 1,000 foot separationbetween aircraft operating at altitudes from 29,000 to 41,000 feet inclusive. Theoperators must also have obtained the airworthiness approvals necessary to flyspecific fleet type aircraft in RVSM designated airspace. Aircraft not complyingwill fly below RVSM airspace. As of January 20, 2005 RVSM airspace nowcovers all of North America, Canada and Mexico.

This reduced separation provides an additional six flight levels and increasedairspace capacity.

Requirements

RVSM maintenance program is FAA governed and altimetry errors mustbe reported to the agency within 96 hours stating the irregularity and correctiveaction. Height keeping errors are monitored by the Aircraft Engineeringdepartment and the AD/Regulatory Compliance group. These two groupsmonitor aircraft logbook discrepancies for height keeping errors.

In order to qualify for RVSM, certain equipment must be installed on the aircraft.

• There must be at least two independent altitude measuring systems (the Captain and First Officers primary altimeters satisfy this requirement).

• There must be at least one altitude reporting transponder. • An altitude alerting system. • An automatic altitude control system.

Manual System

Refer to MEL for each aircraft type for RVSM compliance when deactivating anysystem that directly affects the RVSM airworthiness.

Maintenance of RVSM components is also a critical item. The aircraft illustratedparts catalog (IPC) may denote RVSM-Critical components versus a standardaircraft configuration.

The aircraft structural repair manual (SRM) will contain specific RVSMrequirements to ensure proper maintenance of airframe geometry relative torepairs or alterations made in defined windows surrounding pitot/static probes,static ports, and AOA sensors.

The SRM will provide the limits for the following:

• Skin waviness tolerances • Aerodynamic smoothness tolerances • RVSM critical area dimensions • Static port height tolerances • Pitot tube and combination probe alignment tolerances • Fastener height tolerances • Bulge and skin contour limits in RVSM critical areas • Repair requirements in RVM areas



REDUCED VERTICAL SEPARATION MINIMA



ANTENNA LOCATIONS

Weather radar antenna

The nose radome area contains the weather radar antenna (flat plate).

Glide slope antenna

Left, right, & centerTwo dual-element antennas are installed on the forward pressure bulkhead in the nose radome area

Localizer antennas

Left, right, & centerTwo dual-element antennas are installed on the forward pressure bulkhead in the nose radome area

ATC (Air Traffic Control) antennas

Left & rightTwo blade antennas are installed, one on upper and one on lower forward fuselage. (mode S)

DME (Distance Measuring Equipment) antennas

Left & rightTwo blade antennas are installed on the lower mid fuselage

Radio altimeter antennas

Left, right, & centerSix surface mounted antennas are installed on the lower mid fuselage; 3 transmit and 3 receive antennas

Marker beacon antenna

The marker beacon antenna is installed on the lower mid fuselage

TCAS (Traffic Collision Avoidance System) antennas

Two directional TCAS antennas are installed, one on the top and one on the bottom of the forward fuselage



ANTENNA LOCATIONS



RADIO ALTIMETER - INTRODUCTION

General

The Radio Altimeter System provides accurate terrain clearance altitude information, displayed in the flight compartment, for use by the flight crew. It also provides input to interfacing systems where radio altitude is used in various computations or for the establishment of flight conditions required for warning annunciation’s.

The Radio Altimeter System consists of 3 identical Radio Altimeter Receiver / Transmitter (R/T) units with their associated equipment. All 3 R/T units operate simultaneously, independently from one another. The radio altitude is computed from the time interval a transmitted rf signal needs to travel to the ground and return to the airplane after reflection from the ground. The radio altimeter system operates at altitudes up to 2500 feet and is primarily used in approach, landing and take-off phases of flight. The system operates in the C-band, with a center frequency of 4300 MHz.

Inputs

Each R/T unit, located on the E5-1 rack in the mid-equipment center, transmits RF signals to the ground through a dedicated transmitter antenna. The reflected RF signals are received by a dedicated receiver antenna and routed to the R/T units for altitude computation. All 3 transmitter antennas and all 3 receiver antennas are flush mounted on the forward bottom of the fuselage. Each R/T unit receives a discrete from an air/ground relay, to separate flight segments in the fault memory and to inhibit recording on the ground.

Outputs

Radio altitude output from the R/T units is transmitted to the captain's and first officer's radio altimeter indicators and both EFIS system EADI's for display, as well as to the using systems: autopilot flight director system (AFDS), EICAS, ground proximity warning system (GPWS) and the central warning system.

Circuit breakers for all three radio altimeter systems are located on the P11 panel.



RADIO ALTIMETER - INTRODUCTION



RADIO ALTIMETER COMPONENT LOCATIONS

Radio Altimeter System Components

The radio altimeter system comprises the following components that are located as follows:

• Left, center and right radio altimeter circuit breakers - located on overhead circuit breaker panel (P11).

• Transmitter antennas - located on bottom of the fuselage at station 577: LBL 14.1, BLO, RBL 14.1.

• Receiver antennas - located on bottom of the fuselage at station 621: LBL 14.1, BLO, RBL 14.1.

• Left, center and right receiver/transmitter units - located on rack E5-1 in the mid equipment center.

Interfacing System Components

The following interfacing systems components are associated with the radio altimeter system and are located in the flight compartment as indicated:

• Left and right EFIS control panels - located on quadrant stand P10. • Left EADI - located on captain's instrument panel P1. • Right EADI - located on first officer's instrument panel P3.

Air/ground relay K124 - for left radio altimeter receiver/transmitter, located in left miscellaneous electronic equipment panel (P36).

Air/ground relay K293 - for center radio altimeter receiver/transmitter, located in right miscellaneous electronic equipment panel (P37).

Air/ground relay K214 - for right radio altimeter receiver/transmitter located in right miscellaneous electronic equipment panel (P37).



RADIO ALTIMETER COMPONENT LOCATIONS



RADIO ALTIMETER DISPLAYS

General

On the electronic attitude director indicator, the radio altitude and the decision height are displayed in the right-hand top corner.

Radio Altitude and Decision Height Display

Radio altitude is displayed for altitudes between -20 and 2500 feet. The readout is white, in feet. In addition, if the EFIS "ILS" mode is valid and below 200 feet radio altitude, the green rising runway symbol displays radio altitude by moving up toward the fixed airplane symbol until touchdown. For altitudes above 2500 feet, the readout is blank.

The decision height is displayed above the radio altitude display. The readout is in green and consists of the letters "DH" followed by the selected decision height value in feet between 0 and 999 feet radio altitude. If a negative decision height value is selected, the "DH" display is blanked.

Decision Height Alert and Alert Termination

As the airplane descends through the selected decision height value, the radio altitude readout changes from white to yellow, and the green decision height display changes to the large yellow letters "DH". During the first 3 seconds, the letters "DH" blink.

At reset, the display returns to the normal readout: the radio altitude changes back to white, and the yellow "DH" readout is replaced by the green letters "DH" followed by the selected decision height value.

No-Computed-Data (NCD) and Invalid Data

In the event of no-computed-data, the radio altitude readout is replaced by 4 white dashes. The decision height readout for no-computed-data is replaced by the yellow letters "DH" inside a yellow outline box.

Invalid data is indicated by the yellow letters "DH" and "RA", respectively, each inside a yellow outline box.

Decision Height Alert Termination (Reset)

The decision height alert can be terminated automatically or manually. Automatic reset occurs at touchdown, or when the airplane climbs to a height 75 feet above the selected decision height. Manual reset is achieved by actuating the reset push-button switch "RST" on the EFIS control panel.

Radio Altitude Tape

The functions of the radio altitude tape indicator are to indicate radio altitude, to set the decision height (DH) and to display DH" alert.

The radio altitude tape indicates the airplane altitude above ground. The altitude range is from 0 to 2500 feet. The altitude scale is linear from 0 to 500 feet, and logarithmic from 500 to 2500 feet.

Decision Height Selection

A decision height can be selected within the range from 0 to 499 feet by means of the decision height set knob. The knob sets the decision height index to the selected value, which then is displayed on the three-digit decision height display. The decision height selected and the annunciation on the radio altitude indicator is independent of the one selected on the EFIS control panel. The decision height setting from both radio altitude indicators (higher of the two if different) is an input to the ground proximity warning computer (GPWC) for the mode 6 alert "MINIMUMS-MINIMUMS".

Decision Height Alert and Reset

A decision height alert is annunciated by the illumination of the "DH" light/switch whenever the radio altitude is less than the selected decision height value.

The decision height alert can be reset by pushing the "DH" light/switch. It also is reset automatically when the airplane rises to a radio altitude equal to the decision height value + 15 feet.

Fault Annunciation

The failure flag drops in view for invalid data code or functional test code in the sign-status matrix of the received radio altitude word (ARINC 429), for power supply faults or for malfunction of the indicator internal circuits.In the event of no-computed-data, the indicator readout goes off-scale.



RADIO ALTIMETER DISPLAYS



RADIO ALTIMETER SELF TEST

Self-Test Initiation

The manual self-test is initiated by pressing the TEST switch on the transmitter/receiver front panel. For a complete test, the switch must be held down for at least six seconds. Automatic self-tests are performed at power-on and at regular intervals during normal operation.

Manual self-test capability is inhibited in flight.

T/R Unit Status Indicators

Upon self-test initiation all four front panel status indicators illuminate for three seconds. After the three seconds, the status indicators extinguish for a three-second interval. Subsequently, the green LRU STATUS PASS indicator illuminates for proper system operation. In the event of a fault, the respective red fault indicators illuminate, and the green LRU STATUS PASS indicator remains off. Either LED remains on until the TEST switch is released.

A program exists which provides for the red LRU STATUS FAIL light to come on during self-test in the event of past fault occurrence during the last four flights. This program has several options and is reserved for implementation by the airline. Instructions are obtainable from the vendor representative.

EADI Display

While the test switch is depressed, the EADI indicates a radio altitude of 40 +/- 1 1/2 feet.

Automatic Self-Test

No special test indications are associated with automatic self-tests. If no failures are detected during the automatic self-tests, the operation of the system proceeds normally. If a fault is detected, the fault is annunciated by a yellow RA flag on the EADI.



RADIO ALTIMETER SELF TEST



ALTITUDE ALERT - COMPONENT INTERFACES

Purpose

The system advises the pilots when the airplane approaches within 750 feet of a preselected altitude and when the airplane departs a distance greater than 250 feet from a preselected altitude.

AFCS Mode Control Panel

This panel provides the means for the pilots to input a selected altitude into the AFCS and altitude alert system.

Air Data Inertial Reference Units

These units provide barometric altitude reference data to the altitude alert system.

Proximity Switch Electronics Unit

This unit sends a landing gear up/down signal.

Parking Brake Switch

The set or released status signal is provided by this switch.

Visual

The visual indications output by this system are: amber master caution lights, the amber ALT ALERT light, the two white ALT lights on the captain's and F/Os altimeters, the level B message ALTITUDE ALERT on the upper EICAS display.

Aural

A level B caution aural sound (beep-beep-beep) is heard over both aural warning speakers for 0.8 seconds.

General Operation

Following the selection of a desired altitude, various visual indications and aural sounds occur as the airplane approaches and later deviates from that altitude. Altitude alerting occurs at certain specific distances from the selected altitude during the approach mode and deviation mode. The specific distances are described later.



ALTITUDE ALERT - COMPONENT INTERFACES



ALTITUDE ALERT OPERATION

Operational Sequence

When the airplane approaches the selected altitude and is within 900 feet above or below the selected altitude, visual signals are generated by the altitude alert module. The ALT advisory lights on the captain's and first officer's altimeters illuminate. As the airplane continues toward the selected altitude, and passes through 300 feet from the selected altitude, the ALT advisory lights extinguish. As long as the airplane flies within the 300 feet of the selected altitude, no further indications are produced.

If the altitude deviation subsequently exceeds 300 feet, the following aural and visual indications are produced:

• The aural warning speakers sound the level B caution aural • The amber ALT ALERT light and the master caution lights illuminate • The ALTITUDE ALERT caution message is displayed on the EICAS

Display Unit (upper)

When the pilot changes the selected altitude or when the airplane deviates more than 900 ft from the selected altitude, the caution signals are canceled and the microprocessor is reset to the approach mode. The caution signals are inhibited, in flight, when the landing gear is down and locked. This action prevents nuisance caution indications during the approach phase. The caution signals are also inhibited when the airplane reenters the +/-300 foot envelope above or below selected altitude.

If the airplane is on the ground with the parking brake set, the caution signal inhibits are removed so that the altitude alert system can be tested.



ALTITUDE ALERT OPERATION



ALTITUDE ALERT OPERATIONAL CHECKOUT

Test Preparations

The functional test is accomplished on the ground (landing gear down and locked, parking brake set) by using the mode control panel's altitude select knob to simulate an altitude difference in order to check the approach and the deviation modes.

Operational Checkout

To test the system, slowly rotate the altitude select knob away from the airplane baro altitude as seen on the captain's electric altimeter. Then turn the altitude select knob to approach the airplane baro altitude. Monitor correct operation. Continue rotating the altitude select knob so the error reduces to zero and then increases beyond the deviation threshold. Monitor correct operation.

Use the graphic for the appropriate altitude setting and annunciation’s.

To enable the master caution lights during test, remove the EICAS engine shutdown input (reference MM 31-41-00). For the level B message ALTITUDE ALERT, the level B aural is not inhibited.



ALTITUDE ALERT OPERATIONAL CHECKOUT



INTEGRATED STANDBY FLIGHT DISPLAY (ISFD) COMPO-NENT LOCATION

General

The integrated standby flight display (ISFD) is located on the captains P1 Panel.

The ISFD battery charger is located in the E1 rack.



INTEGRATED STANDBY FLIGHT DISPLAY (ISFD) COMPONENT LOCATION



INTEGRATED STANDBY FLIGHT DISPLAY TEST

General

The integrated standby flight display is tested by pushing both the APP and HP/IN switches together to enter the maintenance mode display. Then push the “TEST” key to enter the maintenance testing menu or “Other Data” to enter other elements of the submenus of the ISFD.



INTEGRATED STANDBY FLIGHT DISPLAY TEST



ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)

General

The EADI's and EHSI are used to display flight and navigation information which includes certain ADIRU data. Attitude and ground speed are shown on the EADI. Track, heading and wind are shown on the EHSI.

The on side (or center if selected) ADIRU is always used by the EFIS symbol generator for attitude and heading displays. Normally the FMC (which uses one or three ADIRU’s) is used by the EFIS for track, ground speed and wind. When the FMC calculated data are invalid, the on side or selected ADIRU is used by the EFIS for track, ground speed and wind displays.

FMC Calculations

The FMC uses inputs from the navigation radios and ADIRU’s to independently calculate ground speed, track and wind vector. The FMC uses North velocity, East velocity and heading from the ADIRS, and true airspeed (TAS) from the ADIRUs. It uses latitude and longitude from radio position data.

For use in its calculations, the FMC first determines the total velocity vector which is based, in part, on the average of the ADIRU’s North and East velocities. If a velocity from one IRU differs from the average by more than 20 knots, the FMC ignores that ADIRU, and uses a single ADIRU (on side or center if selected) to calculate the total velocity vector. Ground speed and track are calculated from the total velocity vector.

To calculate the wind vector, the FMC uses its ground speed and track calculations, TAS and heading from the ADIRS. The FMC uses heading from a single ADIRU which corresponds to the autopilot in command. If no autopilot is in command, the left-most available ADIRU is used.

Track

When the airplane is in the air, the EFIS normally uses the FMC calculated value of track for display. If the FMC track is invalid, EFIS will use track from its on side or selected ADIRU. When the airplane is on the ground (ground speed less than 50 knots), EFIS uses heading from the on side or selected ADIRU and displays it as track (track and heading are always the same on the ground).

If the on side ADIRU fails in flight, the other ADIRU’s continue to provide valid data to the FMC, which provides valid track to the EFIS symbol generators. The EFIS will continue to show a valid map display (except the heading bug is missing). If the on side ADIRU fails on the ground, the map display will show the MAP and TRK flags because the EFIS uses the heading from the on side (which is invalid) ADIRU as a substitute for track.

Ground Speed

Normally the EFIS uses the FMC for the ground speed display. If the FMC is not valid the EFIS uses the on side or selected ADIRU for the ground speed display.

Wind

Normally, the EFIS uses the FMC for the wind direction and speed display. If the FMC is not valid the EFIS uses the on side or selected ADIRU.

Heading and Attitude

The EFIS always uses the on side or selected ADIRU for heading and pitch and roll attitude display.



ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)



EFIS - INTRODUCTION

General

The flight instrument system provides displays for most of the airplane navigational systems.

Subsystems included in the flight instrument system are:

Electronic Flight Instrument System (EFIS)

EFIS includes the electronic attitude director indicators (EADI’s), electronic horizontal situation indicators (EHSI’s), EFIS symbol generators, and EFIS control panels.

Radio Distance Magnetic Indicators (RDMI’s)

The RDMI’s display airplane heading navigational distance, an d directional bearings.

Vertical Speed Indicators (VSI’s)

For display of vertical climb and descent rates as sensed by the Air Data Inertial Reference System (ADIRU’s).

Instrument Source Select Switches

For switching to alternate navigational sources in case the primary sources fail.



EFIS - INTRODUCTION



EFIS COMPONENT LOCATIONS

EADI’s (2) and EHSI’s (2)

Located directly in front of the captain and first officer on P1 and P3.

Instrument Source Select Switches

Located on the instrument source select panels on the outboard edges of P1 and P3.

EFIS Control Panels

Separate panels for the captain and first officer are located on the left and right side of P10.

The Left, Right and Center EFIS Symbol Generators

Located on equipment rack E1.

Remote Light Sensors

Two forward-sensing sensors are located on the glare shield panel P7. (Additional ambient light sensors are integrally mounted on the front faces of the EADI’s and EHSI’s.)

HDG REF Switch

Located on first officer's instrument panel P3-1.

EFIS SYMBOL GENERATOR

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 display CRTs.

Front Panel

The momentary TEST switch initiates the self-test for checking the symbol generator, display units, and control panel. The momentary RESET switch erases the faults stored in memory. The RESET function is not used on the new-generation symbol generators.



EFIS COMPONENT LOCATIONS



EFIS CONTROL PANEL

Purpose

The EFIS control panel controls displays on the EADI and EHSI, allows selection of decision height, and enables the weather radar system.

Switch Functions

EADI Controls:

• BRT - controls brightness level of EADI display. • DH REF - these LCD’s display the selected decision height. • Decision height set knob - this 24-detent, continuous-turn control knob sets

the decision height. The range for decision height is -20 to +999 feet. At selection below zero feet the DH display on the EADI is removed. Decision height starts at 200 feet, as a baseline, when power is applied, and corrected by turning the DH set knob. Two speeds of response are achieved by software.

• RST - manually resets the decision height circuits after the airplane has passed through decision height. EHSI Controls:

• RANGE - selects the range for the weather radar and navigation data displayed on the EHSI.

• TFC - enables TCAS traffic data on the EHSI in MAP, VOR, or ILS modes. • Mode select switch - selects mode of data on the EHSI display. The

modes display 70 degrees arc, with the airplane symbol at the bottom of the display on all modes except the PLAN mode. The CTR (center) MAP switch allows the selection of a center map display as well as a full ILS or full VOR display.

Note: In the PLAN mode, actuation of CTR map switch is mechanically inhibited.

• 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 - this push-on/push-off switch turns on the WXR XCVR and enables the display of weather radar information on the EHSI during the MAP, VOR, or ILS modes. No weather radar data is displayed during PLAN mode. The white band around the rim is visible only in the OFF position.

• MAP display switches - during MAP - mode, 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. The white band around the rim of each cap is visible only in the OFF position.

• NAVAID - VOR, VORTAC, etc. • ARPT - airports • RTE DATA - waypoint altitude and estimated time-of-arrival • WPT - waypoints not in the selected flight plan



EFIS CONTROL PANEL



INSTRUMENT SOURCE SELECT PANELS

Purpose

These two panels allow the pilots to connect to their alternate data sources. The captain and first officer can make selections independently of the other. The ALTN switch illuminates when the alternate source has been selected.

Switch Functions

• FLT DIR Switch - This switch connects the left, center, or right flight control computer to the flight director portion of display on the EADI.

• FMC Switch - This switch selects the left or right flight management computer (FMC), or the on side control display unit (CDU), as the source of navigation and flight parameters for the EHSI display. It is also used to select the source for display on the on side FMC CDU. The switch on the Captain's side only is also used to determine which FMC is the source of the VOR/DME autotune frequency. When FMC-R or CDU-L are selected, the right FMC is the autotune source.

Note: Normally, the CDU-L position is not selected unless both FMC’s are faulty.

• EFI Switch - The EFI switch determines if the on side (normal) or the center (alternate) symbol generator supplies the video presentation on the EADI and EHSI. The captain's and first officer's EFI switches are interlocked electrically such that if both are using the ALTN position, the captain's EFIS control panel and instrument source select switches have control of the center EFIS symbol generator, and the INSTR switch level B EICAS message is initiated.

• IRS Switch - This switch determines which IRU provides data to the on side EFIS symbol generators, and VSI’s, the offside RMI’s, weather radar transceiver(s), the digital flight data acquisition unit (captain's switch only), and the antiskid/autobrake system. The right IRS instrument source select switch also determines if the center IRU receives air data inputs from the left or right air data computer. The on side IRU is normal; the center IRU is alternate.

AIR DATA Switch - Each AIR DATA switch selects air data inputs to the on side or center EFIS symbol generators, on side mach/airspeed indicator, on side electric altimeter, ATC transponder and inertial reference unit. The left switch also selects the altitude source to the altitude alert module and flight recorder.



INSTRUMENT SOURCE SELECT PANELS



OPERATION - NORMAL IR DATA SWITCHING

Instrument Source Select Panel

ADIRS source select switches allow the captain (P1-1) and first officer (P3-3) to switch between the normal (on side ADIRU) and alternate (center IRU) source of ADIRU data.

Switch push-button illuminates white when switch is in ALTN position.



OPERATION - NORMAL IR DATA SWITCHING



EFIS EADI AND EHSI DISPLAYS

EADI display modes

The EADI has only one normal display mode in which the airplane attitude and flight director commands are shown.

EHSI display modes

The EHSI display depends on the position of the mode select switch on the EFIS control panel.

• "PLAN" mode: This mode is generally used prior to flight to set-up the flight plan. The display is oriented north-up.

• "MAP" mode: This FMS mode is used during flight to monitor the airplane's position along the flight plan selected and stored in memory. The display is oriented magnetic track up. Weather radar data can also be displayed in the "map" mode.

"VOR" and "ILS" modes: The two "VOR" modes ("FULL VOR" & "EXP VOR") are used while flying a VOR radial, and the two "ILS" modes ("FULL ILS" & "EXP ILS") display localizer and glideslope deviations during landing. The two expanded ("EXP") modes display only the horizontal situation forward of the

airplane (70o arc), while the two "FULL" modes display a full 360o compass rose. Weather radar data can be displayed only in the "EXP VOR" and the "EXP ILS" modes.



EFIS EADI AND EHSI DISPLAYS



SELF- TEST

General

Self-test patterns are displayed on both EFIS display units when the self-test switch on the symbol generator or the annunciator "test" switch on the P5 panel is pressed. An air/ground relay prevents the EFIS self-test from the P5 panel switch when the airplane is in the air. The P5 panel "test" switch is latching on the -232 airplanes, and it is momentary on the -332 airplanes. In either case be sure to hold the switch for a minimum of 3 seconds or until the test patterns appear. The switch must be held on the -332 airplanes in order to keep the test patterns displayed.

EADI Self-Test Display

The "TEST" message indicates an "OK" condition. In a "FAIL" condition, the faulty LRU is identified by a two-letter code: control panel ("CP"), EADI display unit ("DU"), and symbol generator ("SG") in order of priority. Symbol colors are the same as for normal operation.

EHSI Self-Test Display

The EHSI test pattern displayed during an EFIS self-test depends upon the selected EFIS mode. "test" messages and IRU identification codes are identical to those for the EADI. Certain respective symbols appear in the test pattern only if the appropriate EFIS control panel "map" background switches are actuated ("NAV AID", "ARPT", "RTE DATA" & "WPT"). The weather radar three-sector raster display (red, yellow, green) appears only in "MAP", "EXP VOR" or "EXP ILS" modes.

The P5 panel "TEST" switch will test all three symbol generators simultaneously.



SELF - TEST



VERTICAL SPEED INDICTATORS OVERVIEW

General

The normal data source for the left VSI is the left IRU and left ADC, and the center IRU is the back-up source. Pressing the left "IRS" instrument source select switch to the "ALTN" position switches the data source from the left to the center IRU. The data sources for the right VSI are the right ADIRU or center ADIRU, and are selected by the right "IRS" instrument source select switch.

The VSI fault flag is displayed for faults in the VSI, ADIRU. This is because the vertical speed indicators use the ADIRU as a data source.



VERTICAL SPEED INDICTATORS OVERVIEW



MAGNETIC COMPASS LOCATION

General Component Locations

A standard magnetic compass is mounted in the flight compartment area as shown. This instrument is used as a backup magnetic heading reference.

Mechanical

The compass is 2 5/8 inches in diameter. It is mounted under the overhead P5 panel with a non-ferrous bracket. It is a sealed, liquid filled unit, with a circular indicator card that has two parallel and horizontal magnets attached and free to rotate and tilt as the airplane banks. The liquid medium dampens rapid movements and oscillations.

The front panel has E-W and N-S compensators which are used to correct for both magnetic deviations generated by airplane components and electrical currents in local wiring. A card mounted below the compass is used to record small deviation errors that can not be removed by the compensators.

Electrical

The rotating compass card is lighted by a 5 volt ac bulb. Brightness is controlled by the PANEL lighting control on the left side of the glare shield panel (P7). Power is received from the 115 volt AC STBY BUS via the STBY INSTR LTS circuit breaker and stepped down to 5 volt ac in the pilots' center instrument panel dimmer control station. To change the bulb, loosen and remove the lamp holder assembly.



MAGNETIC COMPASS LOCATION



VHF OMNIDIRECTIONAL RANGE SYSTEM - INTRODUCTION

System Description

The VOR system is made up of two VOR receivers, two VOR control panels and one dual VOR antenna. The VOR receivers have two tuning modes. They can be manually tuned by selecting a frequency on the VOR control panel. They can also be tuned automatically by the Flight Management Computer (FMC). The tuning mode is set according to the EFIS (Electronic Flight Instrument System) mode, selected on the EFIS control panel. When VOR or ILS EFIS modes are selected, a discrete from the VOR control panel tells the VOR receiver to use the frequency commands from the VOR control panel. In other EFIS modes the VOR receiver uses frequency commands from the FMC.

System Operations

The VOR ground station transmits direction-code rf. The receivers extract bearing information and provide data outputs to the RDMI’s or RMI’s, the FMC’s, and the EFIS symbol generators. Audio outputs to the flight interphone system provide station identification.

Flight Deck Display

VOR omni-bearing is sent to the FMC for navigation use (position update). Omni-bearing is also sent to the RDMI’s for direct display of magnetic heading reference omni-bearing to the tuned VOR station. Selected course (selected by the pilot on the VOR control panel), and omni-bearing, are sent to the EFIS symbol generators for display of VOR data on the EHSI (Electronic Horizontal Situation Indicator). The EHSI shows selected course, deviation between selected course and omni-bearing, and a TO/FROM indication.



VHF OMNIDIRECTIONAL RANGE SYSTEM - INTRODUCTION



VOR COMPONENT LOCATIONS

VOR Control Panels

Captain's VOR control panel (CP) on left side of the center glare shield (P55).

First officer's VOR CP on right side of the center glare shield (P55).

VOR Receivers

Left VOR receiver on E2-2.

Right VOR receiver on E2-3.

VOR Antenna

Located under vertical stabilizer fin cap.



VOR COMPONENT LOCATIONS



VOR CONTROL PANEL

Purpose

The VOR control panel provides manual VOR frequency and course selection and display. Manual VOR frequency is also provided to the related DME frequency for paired tuning.

Front Panel Controls

Frequency for manual tuning is selected by the two left concentric knobs. The left window shows a five-digit liquid crystal display (LCD), to indicate the VOR receiver tuned frequency for both manual and auto tuning modes. The reading for 100 MHz is fixed. The tens and units MHz is selected by the outer knob, and the last two digits is selected by the inner knob.

The tuning knob annunciator indicates the tuning mode of the VOR system and allows the pilot to manually override the automatic tuning mode.

The course select knob is a ten-turn control knob. The course selected is displayed on the three-digit liquid crystal display (LCD) window above the course select knob.



VOR CONTROL PANEL



VOR TUNING - CONTROLS AND DISPLAYS

General

The VOR frequency can be tuned manually by the pilot, or it can be tuned automatically by the flight management computer. It can also be tuned remotely from the FMC control display unit (CDU).

Manual Tuning

Initialization - On the EFIS control panel, set the display select switch to VOR. The AUTO/MAN override switch on the on side VOR control panel indicates MAN.

On the VOR control panel, select the VOR frequency and select the VOR course. Both selections will be displayed on the VOR control panel.

Data Verification - On the RDMI, set the display select switch to VOR and verify that bearing indicator points to the VOR station.

On the EHSI, verify that VOR source annunciator, selected course, and VOR deviation bar are displayed.

On the FMC CDU, press the PROG key and verify that tuned frequency on line 5L or 5R agrees with frequency on the VOR control panel.

Automatic Tuning

Initialization - On the EFIS control panel, set the display select switch to MAP or PLAN.

On the VOR control panel, the AUTO/MAN manual override switch/light indicates AUTO.

Data Verification - On the RDMI, set the display switch to VOR and verify that the bearing indicator points to the VOR station.

On the FMC CDU, press the PROG key and verify that the tuned frequency on line 5L or 5R agrees with frequency on VOR control panel.

VOR Remote Tune - The FMC selected VOR station can be changed while in the automatic tuning mode, using the FMC CDU.

On the FMC CDU, press the PROG key. Enter the VOR station identifier (if stored in FMC data base) or frequency into the scratchpad. Line select the scratchpad to line 5L or 5R to remote tune the left or right VOR, respectively.

To return the remote tune mode to the autotune mode, type A in the scratchpad and line select it to the remote tune side.



VOR TUNING - CONTROLS AND DISPLAYS



VOR DISPLAYS

Electronic Horizontal Situation Indicator

Normal Displays - During a VOR mode, the EHSI displays data as illustrated in the graphic. This includes VOR selected course, course deviation, a to/from indicator, and a data source indicator (VOR-L or VOR-R).

NCD Display - If the omni-bearing data word is NCD, the deviation bar and to/from indicator will be removed. The scale and course pointer is retained.

Invalid Display - If the omni-bearing data word is invalid, the deviation bar and scale is removed and the yellow VOR flag is shown. The course pointer is retained. If only the selected course data word is invalid, the course pointer, deviation

Radio Distance Magnetic Indicator

The RDMI’s display VOR and ADF bearing on the two rotating bearing pointers. The type of bearing displayed is selected by the bearing pointer source control knobs. The pointer (bearing) flags drop into view for invalid or no computed data.

The VOR bearing flags will also be in view when ever magnetic heading is failed or NCD. The bar, and to/from indicator are removed. The scale is retained.



VOR DISPLAYS



VOR SELF-TEST (MAIN EQUIPMENT CENTER)

Functional Test

The self-test sequence is executed by pressing the test switch on the front panel of the VOR receiver for approximately 10 seconds. The results of the test appear on the LED status lights until the test switch is released or until the self-test automatically times out after approximately 9 +/-1 second (whichever occurs first).

Front Panel Features

Test switch - this switch provides the self-test for both the VOR receiver circuits and the marker beacon receiver module.

VOR LED (red) - indicates a detected failure in the VOR receiver.

VOR LED (green) - indicates no detected faults in the VOR receiver. Red LED’s are off.

MKR LED (red) - indicates a detected failure in the marker beacon module, irrespective of whether the module is in the L VOR or R VOR receiver.

DATA IN LED (red) - indicates invalid data from the selected tuning source. This source can be the VOR control panel, L FMC, or R FMC.

The status indicating LED’s illuminate only during manual self-test.



VOR SELF-TEST (MAIN EQUIPMENT CENTER)



DME - INTRODUCTION

General

The distance measuring equipment (DME) measures the slant-range distance from the airplane to a selected ground station and provides continuous distance information to the flight management computers for high accuracy position fixing and simultaneously provides data available for DME distance displays. An audio output provides station identification.

The DME system measures distance by transmitting a pulse pair signal to a ground station and counting the time it takes to receive a reply signal. The distance is then the propagation velocity of the pulse pairs multiplied by the time divided by 2. The factor of 2 is necessary because the pulse pair must traverse the distance twice, first down to the station, and then back to the interrogator.

Frequencies

The distance measuring equipment (DME) ground stations are co-located with ILS or VOR ground stations. Even though the DME's operating frequencies are in the UHF band, the DME interrogators are tuned by providing them with the VHF ILS or VOR frequency of the station with which they are co-located. The interrogator then translates the VHF frequency into the corresponding UHF frequency of the DME ground station.

System Controls

The system mode control is from the EFIS control panel. If the selected EFIS mode is ILS, the DME frequency is paired with the selected ILS frequency. If the selected EFIS mode is VOR, the DME frequency is paired with the selected VOR frequency. The interrogator is said to be manually tuned in the ILS or VOR modes. In the EFIS map or plan modes (frequency scanning mode), the DME frequency is controlled by the flight management computer (FMC). The FMC directs the tuning of up to five DME ground stations. The interrogator is said to be autotuned in the map or plan modes.

Outputs

The DME distance to the station is displayed on both RDMI’s and the EHSI’s. Distance data is also supplied to the FMC’s and the digital flight data acquisition unit (DFDAU). Audio is supplied to the flight interphone system for station identification.

Purpose

The DME interrogator transmits a pulse-pair interrogation signal, receives the reply signal, computes the slant range to the tuned station, and provides distance output data.

Front Face

The test push-button switch initiates the interrogator self-test sequence. Test sequence indications are output to interfacing systems and DME status is indicated on the front panel LED’s. The status indicating LED’s illuminate only during manual self-test. The red R/T, ANT, and DATA IN LED’s indicate faults in the interrogator, DME ANT DC continuity, and frequency tuning source, respectively. The green R/T LED indicates a no-fault condition (no red LED’s illuminated).

Note: The antenna monitoring is a customer option. Fault Memory (Customer Option)

A non-volatile fault memory is included (customer option) which stores up to 13 faults per flight for the last 64 flights. If two similar faults are stored in the four most recent flights, the front panel red R/T LED will light during manual self-test. Flight legs are determined by a discrete from an air/ground relay. Fault memory data is fed to a rear panel ATE connector upon a dump request from shop level test equipment. On the ground, fault recording is inhibited except during manual self-test.



DME - INTRODUCTION



DME COMPONENT LOCATIONS

DME Interrogators (Left System and Right System)

On main E/E rack in main equipment center (E2-2 and E2-3).

DME Antennas (Right and Left)

On bottom of fuselage, left at station 555, right at station 665.

Radio Distance Magnetic Indicator (RDMI)

Left on captain's instrument panel (P1)

Right on first officer's instrument panel (P3)



DME COMPONENT LOCATIONS



DME CONTROLS AND DISPLAYS

DME Controls

The DME interrogator tuning mode is selected on the on side EFIS control panel. The mode selected determines whether the DME is manually tuned or automatically tuned by the FMC. If an auto-tune mode is selected, the on side VOR control panel gives the pilot manual tuning override capability.

EFIS Control Panel - When FULL VOR or EXP VOR is selected on the EFIS control panel, the DME is tuned by a paired VOR frequency. This frequency is manually selected on the VOR control panel, and routed to the DME.

When FULL ILS or EXP ILS is selected on the EFIS control panel, the DME is tuned by a paired ILS frequency. This frequency is manually selected on the ILS control panel and routed to the VOR control panel. In the VOR control panel, the ILS frequency is routed through a relay to the DME. The relay sends the ILS or VOR frequency to the DME depending on the mode selected.

When map or plan mode is selected on the EFIS control panel, the DME interrogator is automatically tuned by the flight management computer. The FMC directs the interrogator to tune up to five DME stations. The AUTO part of the VOR/DME switch-light on the VOR control panel is lighted. Pressing this switch light overrides the auto-tune mode and causes the DME to revert back to manual tuning by the selected VOR frequency. The MAN part of the VOR/DME switch-light will be lighted. Pressing the switch-light a second time brings back the auto tune mode.

VOR Control Panel - The VOR frequency, which is manually selected with knobs on the VOR control panel, is sent to the DME when the VOR modes are selected. The selected frequency is shown on an LCD display. The MAN part of the VOR/DME switch-light will be lighted white.

Each VOR station has a collocated DME station. The VOR frequency and the DME frequency are paired. The DME interrogator accepts the VOR frequency and uses a memory look-up table to determine the actual DME frequency to be tuned.

ILS Control Panel - The ILS frequency, which is manually selected with knobs on the ILS control panel, is sent to the DME interrogator by way of the VOR control panel when the ILS modes are selected. The selected frequency is shown on an LCD display. The MAN part of the VOR/DME switch-light will be lighted.

The DME interrogator accepts the ILS frequency and uses a memory look-up table to determine the actual DME frequency to be tuned.

DME Distance Displays

Radio Distance Magnetic Indicator (RDMI) - The RDMI displays DME/VOR distance between 0 to 799.9 to the tenth of a nautical mile. The DME display shows blank for DME faults and dashes for DME no-computed-data or DME distance out of range. If the EFIS control panel mode is ILS, the DME/ILS distance includes an L in the most significant digit position, and the maximum display value is 99.9 nm instead of 799.9 nm. If the EFIS control panel mode is map or plan, the DME distance is provided by the FMC.

Electronic Horizontal Situation Indicator (EHSI) - The EHSI displays DME distance during VOR and ILS modes. The annunciation DME is also displayed. During map and plan modes, the distance displayed is distance-to-go from the flight management computer and is not from DME, although DME may be providing distance information to the flight management computer. The DME distance display is to the nearest nautical mile if distance is greater than or equal to 100 nm, and to the nearest tenth of a nautical mile for distance less than 100 nm. DME faults cause the display to be blank, and DME no-computed-data causes a display of four dashes.



DME CONTROLS AND DISPLAYS



DME DISTANCE DISPLAY

DME Distance on EHSI

DME distance is shown on the EHSI while VOR or ILS modes are selected on the EFIS control panel.

Note: When the MAP or PLAN modes are selected on the EFIS control panel, DME distance is replaced by distance to the next waypoint.

DME Distance Display on RDMI

DME distance is shown on the RDMI in all EFIS modes.



DME DISTANCE DISPLAY



DME SYSTEM - INTERFACES

System Configuration

Two DME Interrogators and two DME Antennas, one for each interrogator, are installed in the airplane.

Frequency Control

The pilot determines the tuning mode (manual or automatic) of the DME interrogator by his mode selection on the on side EFIS control panel. Also, the on side VOR control panel gives the pilot manual tuning override capability.

EFIS VOR Mode - With the VOR mode selected on the EFIS control panel, the manual relay is energized and the ILS relay in the VOR control panel is relaxed by the tuning logic. The tuning source select discrete to the DME interrogator is a logic one (open), so the DME interrogator accepts the manual tune frequency selected by the pilot on the VOR control panel.

EFIS ILS Mode - With the ILS mode selected on the EFIS control panel, the manual relay and the ILS relay in the VOR control panel are energized. The tuning source select discrete to the DME interrogator is a logic one (open), so the DME interrogator accepts the manual tune frequency selected by the pilot on the ILS control panel.

With the MAP or PLAN mode selected on the EFIS control panel, both the manual relay and the ILS relay in the VOR control panel are relaxed. The tuning source select discrete to the DME interrogator and the FMC’s is a logic zero (ground) so the DME interrogator accepts the auto-tune frequency output by the FMC. The FMC’s output a left or right DME scan discrete (ground) which energizes K1304, scanning right RDMI relay, or K1303, scanning DME left RDMI relay.

The left FMC is the normal auto-tune frequency source for both DME interrogators. If the captain selects CDU-L or FMC-R on his FMC alternate source select switch, the left and right FMC tuning relays (K757 and K758) energize, and the right FMC replaces the left FMC as the auto-tune frequency source. The FMC’s direct the tuning of up to five DME ground stations.

Manual Tuning Override Capability - A latch within the tuning logic in the VOR control panel is enabled with Map or Plan selected on the on-side EFIS control panel. Otherwise it is disabled.

With MAP or PLAN mode selected on the EFIS control panel, the manual relay is relaxed and the ILS relay is relaxed. If the AUTO/MAN switch on the VOR control panel is pressed, the manual relay is latched in the energized state. The tuning source select discrete to the DME interrogator is a logic one (open) so the DME interrogator accepts the manual tune frequency selected by the pilot on the VOR control panel. The manual relay will be held in the energized state until the AUTO/MAN switch is pressed a second time, or VOR or ILS mode discrete from the EFIS control panel change state (open to ground).

Suppression

Each DME interrogator generates a suppression pulse for use internally and for suppressing the receivers of the opposite DME, both ATC transponder’s, and TCAS computer when interrogation pulses are being transmitted. In addition, the DME interrogator accepts suppression pulses to protect its receiver when the other L-band equipment is transmitting.

DME Distance Output

The DME interrogator determines the slant range distance to the ground station it is tuned to by measuring precisely the amount of time that has elapsed between transmission of an interrogation rf pulse pair and reception of the reply rf pulse pair.

DME distance data is sent from each interrogator to both FMC’s for navigation position fixing, and is sent for display to both RDMI’s, and to the on side and center EFIS symbol generators for display on the EHSI’s. The FMC’s provide distance data to the RDMI’s in EFIS MAP or PLAN mode.

DME Audio Output

Each DME ground station periodically transmits a 1350 Hz Morse coded identification signal. This signal is decoded in the DME Interrogator and routed through the on side VOR control panel to the audio selector panels.



DME SYSTEM - INTERFACES



DME SYSTEM TEST

DME Self-Test

A manual self test is initiated by pressing the test switch on the front panel of the interrogator and holding it for the duration of the test. This will give the flight deck indications as well as the DME status indications on the front panel.



DME SYSTEM TEST



GLOBAL POSITIONING SYSTEM - INTRODUCTION

General

The global positioning system (GPS) uses navigation satellites to supply accurate airplane position to airplane systems and to the flight crew.

Abbreviations and Acronyms

• AIL autonomous integrity limit • BITE built-in test equipment • CDU control display unit • D/A digital-to-analog • FMCS flight management computer system • GPS global positioning system • GPWC ground proximity warning computer • HFOM horizontal figure of merit • MMR multi-mode receiver • NCD no computed data • PPS precision positioning service • RAIM receiver autonomous integrity monitor • SPS standard positioning service • UTC universal time (coordinated) • RAIM receiver autonomous integrity monitoring



GLOBAL POSITIONING SYSTEM - INTRODUCTION



GPS THEORY OF OPERATION

GPS Segments

The GPS has three segments:

• Satellite • User • Control

Satellite Segment

The satellite segment is a group of satellites that orbit approximately 10,900 nautical miles above the earth. Each satellite completes an orbit approximatelyonce every 12 hours. There are 24 operational satellites and 3 spares.

The satellites continuously transmit radio signals with navigation data, range code, and the exact time.

User Segment

The user segment consists of the GPS receivers which are in the multi-mode receivers (MMR). They receive satellite signals from their preamplified antennas and use them to calculate the airplane position by distance calculation to all visible satellites.

Control Segment

The control segment has control and monitor stations on earth that continuously monitor and track the satellites. The purpose of the control segment is to do these functions:

• Monitor and correct satellite orbits and clocks • Calculate and format a satellite navigation message • Update the satellite navigation message regularly

The control segment has one master control station and five monitor stations. Three of the monitor stations are also upload stations.

The master control station is in Colorado Springs, CO, USA. The master control station is the operational center of the GPS. The master control station controlsall operations in the control segment. The master control station has an atomic clock. This clock is the reference for the GPS.

The monitor stations track the satellites 24 hours a day. The master control station remotely controls the monitor stations through on-line connections. Themonitor stations are in these locations:

• Ascension island • Colorado Springs • Diego Garcia island • Hawaii • Kawajalein island

To calculate the airplane position (latitude, longitude, and altitude) and the clock bias, the GPS receiver must know the position of at least four satellites. It then measures the distances to all the satellites at the same time, and solves for these four unknowns with four range equations:

• Latitude • Longitude • Altitude • Clock bias

GPS Time

All the satellites synchronize to universal time (coordinated) (UTC). The satellites transmit this time to the GPS receiver. The accuracy of the satellite UTC is approximately 100 nanoseconds. The GPS receiver transmits UTC on an ARINC 429 format.



GPS THEORY OF OPERATION



GPS GENERAL DESCRIPTION

General

The global positioning system (GPS) receives signals from GPS satellites and calculates:

• Latitude • Longitude • Altitude • Ground speed • Universal time (coordinated) (UTC)

There are two GPS antennas. The left and right antennas receive satellite signals and send them to the GPS section of their multi-mode receivers (MMR).

The GPS receivers calculate airplane position and report Universal time coordinated (UTC). Time and position data is used to update the flight management computers. Position only data goes to the enhanced ground proximity warning computer (EGPWC) to compare airplane position with the EGPWC database.

The GPS also provides UTC to the captain and first officer clocks. If this provision is activated, the clocks show GPS time.



GPS GENERAL DESCRIPTION



GPS COMPONENT LOCATIONS

Antenna

The left and right GPS antennas are on top centerline of the fuselage at stations 600 and 622. The antennas are active (amplified) and receive 12v dc power from their onside multimode receiver.

Multimode Receiver

The left multimode receiver (MMR) is on the E1-3 shelf in the main equipment center (MEC). The right MMR is on the E1-5.



GPS COMPONENT LOCATIONS



GPS COMPONENTS (CONT’D)

General

The GPS antennas receive L-band frequency signals and sends them to the multi-mode receivers (MMR).

The GPS antennas use built in amplifiers to reduce RF signal loss. The antenna preamplifiers use 12v dc from the power supply in the MMR.

Purpose

The GPS receiver is an electronic card in the multi- mode receiver (MMR) in the main equipment center. It calculates the satellite range solutions from up to twelve satellites at one time. It processes the range solutions and gives aircraft position to the flight management computers and the enhanced ground proximity warning computer. It also sends universal time coordinated (UTC) to the FMC and the captain and first officer clocks.

Front Panel Features

The multi-mode receiver (MMR) has a test switch and three LED status indicators on the front panel. These indicators come on only when the front panel TEST switch is operated:

• Red/green LRU STATUS light - RED shows the receiver has a fault • GREEN shows the receiver is good • Red/CONTROL FAIL - shows control input faults • Red/ANT FAIL - Red shows that any one of the antennas connected to the MMR has failed

General

The only system tests for the GPS are through the MMR.

Each multi-mode receiver has a functional test button that starts a self-test of the MMR, any control interfaces with it, and the antenna inputs to it.

System Test

To start the self-test sequence, push and release the TEST switch on the MMR front panel:

• For the first two seconds, the LRU SATUS, CONTROL FAIL, and ANT FAIL LEDs are red.

• For the next two seconds, the LRU status LED is green, and the control input and antenna LEDs are red.

• For the next two seconds, (minimum) all LEDs are off. After this, the appropriate LEDs show the system status.

• LRU status pass - The green LRU status LED comes on if no faults sre found during the self-test sequence in either the ILS and GPS.

• LRU status fail - The red LRU status LED comes on if a fault is found during the self-test sequence.

• Control input fail - The red LED comes on if the ILS does not receive tuning information from the ILS control panel.

• Antenna fail - The red LED comes on if any (ILS or GPS) system antenna fails.

Power-up

During power-up, the MMRs do a test of the antennas. At that time, the MMRs verify continuity of the antenna connections. The MMRs do not do a test of the antennas during operation after power-up.



GPS COMPONENT LOCATIONS (CONT’D)



GPS FUNCTIONAL DIAGRAM

Power

The power supply provides various dc voltages from the 115v ac input to the multi-mode receiver (MMR).

Satellite Signal Processing

The low noise amplifier (LNA) receives and amplifies the satellite signals from the GPS antenna. The receiver detects the satellite signal and sends it toan analog-to-digital (A/D) converter. The A/D converter sends the digitized signal to the microprocessor. The microprocessor calculates airplane position, altitude, and other GPS data and sends it to the flight management computers (FMCs) and EICAS computers on ARINC 429 buses.

Inputs

The GPS receivers use position data from the IRS during initialization. The GPS also uses this data in the aided and altitude aided modes.

Outputs

Both GPS receivers send time and position data to both FMC’s.

The GPS also sends position data to the enhanced ground proximity warning computer (EGPWC). The GPWC uses this data in its terrain awareness and terrain clearance floor functions.

There are provisions for GPS data to go digitally to the captain and first officer clocks.



GPS FUNCTIONAL DIAGRAM



GPS OPERATION MODES

GPS Modes of Operation

The GPS receiver operates in these modes:

• Acquisition mode • Navigation mode • Altitude aided mode • Aided mode

Acquisition Mode

The GPS receiver looks for and locks on to the satellite signals. The GPS receiver must find at least 4 satellites before it starts to calculate GPS data.

The GPS receiver can accept inertial reference data to calculate which satellites are available at the present airplane position to help to acquire satellites. It canalso reference time and date from the captain and first officer clocks.

If the IRS data is not available, the GPS receiver can still acquire satellites signals. However, it takes longer because it has to look for all the satellites.When the GPS receiver finds the satellites, it calculates which it can use.

The GPS receiver takes approximately 75 seconds to acquire satellite signals when IRS data is available. It can take 4 to 10 minutes to acquire satellites when IRS data is not available.

Navigation Mode

The GPS receiver enters the navigation mode after it acquires and locks on to at least 4 satellites. When the GPS receiver is in the navigation mode, itcalculates GPS data. GPS receiver output goes no computed data (NCD) when the accuracy is not within 16 nautical miles of the actual position.

Altitude Aided Mode

With four satellites available, the GPS receiver stores the difference between the IRS inertial altitude and altitude.

The GPS receiver stores the difference between inertial and GPS altitude so that it can estimate the GPS altitude when only three satellites are available.In the altitude aided mode, the GPS receiver uses the airplane altitude from the IRS and the length of the earth radius as the fourth range.

The GPS receiver enters the altitude aided mode only after these three conditions are true:

• The GPS receiver was in the navigation mode • There are only three satellites available with good geometry for position

fixes • The GPS receiver stored the difference between inertial and GPS altitude

in memory

The GPS receiver starts normal operation again when a fourth satellite comes into view.

Aided Mode

The GPS receiver enters the aided mode during short periods (less than 30 seconds) of bad satellite coverage. An example of bad satellite coverage is poor satellite geometry when at least four satellites are available but they are not spread out far enough so the GPS receiver can make an accurate position fix.

In the aided mode, the GPS receiver receives altitude, heading, and speed from the inertial reference system. The GPS receiver uses this data to go back quickly to the navigation mode when there is good satellite coverage again. The GPS receiver output is NCD in the aided mode.

Autonomous Integrity Limit

The GPS receiver has a receiver autonomous integrity monitor (RAIM) function. The RAIM monitors the status of the satellites that the GPS receiver uses forcalculations. The output of the RAIM function is an estimate of the GPS position error.



GPS OPERATION MODES



GPS POSTITION INITIALIZATION PAGE

General

The flight management computer shows position initialization and position reference pages on the multi-purpose control display unit (MCDU).

The multi-mode receivers (MMRs) send GPS data to the FMC. GPS data shows on the MCDU. The position initialization page shows the GPS position and GPS time.

The flight crew can use the GPS position to initialize the inertial reference system (IRS) position. The GPS universal time (coordinated) (UTC) shows on the MCDU when the GPS time is valid.



GPS POSITION INITIALIZATION PAGE



GPS POSITION REFERENCE PAGE WITH BEARING DISTANCE

General

Position reference page two, line 1L shows the FMC airplane position in large font, and shows the sensor used to calculate the FMC position in small font.

FMC can use GPS position data to calculate the airplane position. It can also use these other navigation systems to calculate airplane position:

• Inertial Reference SYSTEM (IRS) • Distance Measuring Equipment (DME) • VHF Omnidirectional Range (VOR) • ILS Localizer (LOC)

The FMC calculates the accuracy of the position data from each navigation system. These calculations show on POS REF pages 2, 3, and 4 (of 4).

The left FMC uses the left MMR and the right FMC uses the right MMR. If the onside MMR fails, the FMC’s use the other MMR.

Use line select key 6R to toggle between a sensor position in Latitude and Longitude, and a sensor position in bearing and distance relative to the FMC calculated position.



GPS POSITION REFERENCE PAGE WITH BEARING DISTANCE



RADIO DISTANCE MAGNETIC INDICATOR (RDMI)

Purpose

The RDMI’s display magnetic heading, VOR or ADF bearing, and DME distance.

Component Locations

Two RDMI’s are provided; one for the captain on P1, and one for the first officer on P3.

Normal Displays

The compass card displays heading from the offside pilot's selected IRU. Magnetic heading is displayed between 73 degrees N and 60 degrees S latitude with the HDG REF switch in NORM. At greater latitudes, with HDG REF in NORM, magnetic heading is tagged no-computed-data; consequently, the RDMI display the heading card failure flag. In TRUE position, the HDG REF switch reprograms the RDMI’s to use true heading to drive the compass card.

The bearing pointers are driven by a data source selected by the ADF/VOR switch. In VOR mode, the pointer is driven by the respective VOR receiver.

In ADF mode, the pointer is driven by an ADF receiver, or goes into the no-computed-data mode (fault flag) if no ADF receiver is installed for that pointer.

DME distance is displayed in both ADF and VOR modes to the nearest tenth of a nautical mile. In the EFIS ILS mode, the DME distance is preceded by the letter L.

Fault Displays

Heading Flags - The heading flag is displayed on no-computed or invalid data from the IRU supplying heading information, for RDMI circuitry failure, or at latitudes greater than 73 degrees N or 60 degrees S and the HDG REF switch in NORM.

Bearing Flags - VOR/ADF - displayed for a data source failure, no-computed-data, or an RDMI circuitry failure.

VOR flag - also displayed when the heading flag is in view or when the HDG REF switch is in TRUE.

Distance Annunciator - Blank - invalid data from the source or a tuning discrepancy between received data and EFIS control panel selection or RDMI circuitry failure. Dashes - no computed data.



RADIO DISTANCE MAGNETIC INDICATOR (RDMI)



RDMI DIAGRAM

RDMI Data Sources

The various RDMI displays on the L-RDMI present information from the following data sources:

• Magnetic/true heading - (Normal: R-IRU; Alternate: C-IRU). • Heading reference select switch - Override switch forces true heading

display. • Left bearing needle - L-VOR; L-ADF. • Right bearing needle - R-VOR; R-ADF (ER Only). • Left distance readout - L DME; L OR R FMC. • Right distance readout - R DME; L OR R FMC.

Data Source Selections

The displays for the R-RDMI are the same except that the normal heading data source is the L-IRU.

The left distance displays shown at any one time on the RDMI’s depend upon the left EFIS control panel mode (VOR, ILS, MAP, or PLAN mode).

The right distance displays depend upon the right EFIS control panel mode

selection.



RDMI DIAGRAM



ILS - INTRODUCTION

Purpose

The ILS provides airplane position data relative to the glide slope and runway centerline.

The Glideslope and Localizer signals come from ground based transmitters through different antennas. The left, center and right receivers use signals to get glideslope and localizer deviation. The receivers also use Morse code signals from the localizer part of the signal. These are sent to the flight interphone system.



ILS - INTRODUCTION



ILS - GENERAL DESCRIPTION

Purpose


System Description

ILS Control Panel - A single control panel provides frequency tuning and front course runway heading to three receivers simultaneously.

ILS Receivers - Left, right, and center receivers are operating at the same time. Glide slope and Localizer signals are received from ground station transmitters through separate antennas and processed to obtain glide slope and Localizer deviation values. Morse code signals bearing station identification are extracted from the composite localizer transmission and routed to the audio selector panels.

Interface Systems

Automatic Flight Control System - Left, center, and right FCC accept ILS digital data from the respective receiver to generate guidance commands during AFCS operation.

Flight Management System - Left and right FMC’s accept ILS digital data from the on side receiver for the purpose of position updating during final approach.

Electronic Flight Instrument Displays - Glide slope and Localizer deviation displays appear on the EHSI and EADI. The EHSI also displays ILS frequency and runway heading. The left ILS provides data to the captain's EFIS, the right ILS to the first officer's. The center ILS is an alternate data source to either captain's or first officer's EFIS.

Flight Data Recording System - ILS data are processed in the digital flight data acquisition unit, and routed to the flight data recorder.

Ground Proximity Warning System - The GPWC accepts left ILS digital data to trigger below glide slope approach warnings.

Standby Attitude/ILS Indicator - The standby attitude/ILS indicator displays glide slope and localizer deviation data from the center ILS as a backup indicator.

Flight Interphone System - Station audio identification is provided to the audio selector panels from all ILS receivers.



ILS - GENERAL DESCRIPTION



ILS COMPONENT LOCATIONS


ILS receivers - located in the main equipment center, on rack E1-3 (left ILS), E1-4 (center ILS), and E1-5 (right ILS).

ILS control panel - located on the aft pilot's control stand (P8).

ILS antennas - two dual loop G/S antennas and two dual loop LOC antennas are located on the forward pressure bulkhead (section 41) within the nose radome.

ILS circuit breakers - located on the overhead circuit breaker panel P11-1 (left and center ILS) and P11-4 (right ILS).

Glide slope director element - located in nose radome assembly, positioned horizontally across the center butt line on the inside surface of the radome, about 18 inches from the radome lower edge.



ILS COMPONENT LOCATIONS



ILS CONTROL PANEL

Purpose

The ILS control panel encodes and outputs on ARINC 429 data buses the ILS frequency and the front course runway heading as selected by the pilot input controls. The data buses are triple redundant with the encoded data generated by three electrically isolated sets of switches, encoding electronics and sent to the receivers by three dedicated output connectors.

Power requirements - 115 volts ac, 400 Hz, single-phase, 10.3 w maximum.

Control Panel Features

ILS Frequency Display - Displays the center ILS tune frequency or a park display (five-dashes), indicating that the frequency selector is in off (park) position. The display range is 108.10 to 111.95 MHz, indicating the ILS localizer frequency value. The most significant digit is fixed at 1.

ILS Frequency Selector - Selects one of forty localizer frequencies to tune all ILS receivers. Only the LOC frequency is dialed in; the corresponding G/S frequency is determined automatically by the receiver. The outer knob has 10 positions: Park, 08, 09, 10, 11, and repeats. The inner knob has 10 positions: .10, .15, .30, .35, .50, .55, .70, .75, .90, and .95. Park causes the control panel to send a NCD ILS frequency word to the receivers, which places them in standby mode: LOC and G/S deviation outputs are NCD. Also, a park discrete signal to the EFIS symbol generators cause removal of all ILS indications from EADI.

ILS Front Course Display - Displays the course set by the course selector. It indicates the selected runway heading in degrees (000 to 359).

ILS Front Course Selector - Sets the ILS course in the course indicator and on the EHSI if it's operating in the ILS mode. The outer knob controls the hundredths and tenths position and the inner knob controls the units position.

ILS TEST Switch - Covered at the end of this section.



ILS CONTROL PANEL



ILS SYSTEM DIAGRAM

Purpose


System Description

ILS Control Panel - A single control panel provides frequency tuning and front course runway heading to three receivers simultaneously.

ILS Receivers - Left, right, and center receivers are operating at the same time. Glide slope and Localizer signals are received from ground station transmitters through separate antennas and processed to obtain glide slope and Localizer deviation values. Morse code signals bearing station identification are extracted from the composite localizer transmission and routed to the audio selector panels.

Interface Systems

Automatic Flight Control System - Left, center, and right FCC accept ILS digital data from the respective receiver to generate guidance commands during AFCS operation.

Flight Management System - Left and right FMC’s accept ILS digital data from the on side receiver for the purpose of position updating during final approach.

Electronic Flight Instrument Displays - Glide slope and Localizer deviation displays appear on the EHSI and EADI. The EHSI also displays ILS frequency and runway heading. The left ILS provides data to the captain's EFIS, the right ILS to the first officer's. The center ILS is an alternate data source to either captain's or first officer's EFIS.

Flight Data Recording System - ILS data are processed in the digital flight data acquisition unit, and routed to the flight data recorder.

Ground Proximity Warning System - The GPWC accepts left ILS digital data to trigger below glide slope approach warnings.

Standby Attitude/ILS Indicator - The standby attitude/ILS indicator displays glide slope and localizer deviation data from the center ILS as a backup indicator.

Flight Interphone System - Station audio identification is provided to the audio selector panels from all ILS receivers.



ILS SYSTEM DIAGRAM



ILS THEORY OF OPERATION

Localizer Theory

A localizer signal is transmitted on one of 40 frequencies. The frequency range is 108.10 to 111.95 MHz, on odd tenths.

The localizer beam is aligned with the final approach course to the runway. The localizer transmitter produces two lobes, one on either side of the runway centerline. The left lobe is modulated with a 90-Hz carrier and the right lobe with 150 Hz. The final approach course coincides with the runway centerline, and is the course where the two signals (90 and 150 Hz) are equal.

If the airplane is to the left of localizer centerline, the localizer deviation display moves to the right. This means the runway centerline is to the right. If the airplane is to the right, the localizer deviation display moves to the left. This means the runway centerline is to the left.

The amount of deflection, left or right, depends on the relative strength of the two signals. If the airplane is one degree left or right of course, the receiver moves the deviation bar one dot left or right. If the airplane is two degrees left or right, the receiver makes a two-dot deflection.

Many airports have only one instrument landing system (ILS). An approach to the airport from the opposite direction is called a back course approach. The ground station uses the same localizer transmitter for a back course (B/CRS) approach. The modulation frequencies stay on the same side of the runway. Therefore, the deviation bar would deflect in the opposite direction with the normal final approach course selected. However, EFIS compares the selected approach course with aircraft heading. If there is a difference of more than 90 degrees, it reverses the signal polarity and the indications show correctly. The crew must select B/CRS on the standby attitude/ILS indicator or the integrated standby flight display (ISFD) for that display to show correctly.

Glideslope Theory

The glideslope signal is transmitted on one of 40 frequencies. The range is from 329.15 to 335.0 MHz. The glideslope transmitter is automatically tuned when the crew selects an ILS localizer frequency.

Glideslope signals produce two lobes, one above the other. The upper lobe is modulated with 90 Hz, and the lower lobe with 150 Hz. The glide slope transmits from a point past the runway threshold in the direction of the localizer. It provides a 2.5 to 3 degree glide path where the two audio signals are equal.If the airplane is above the glide path, the glideslope deviation display moves down to show the glideslope centerline is below the airplane. If the airplane is below the glide path, the glideslope deviation display moves up to show the glideslope centerline is above the airplane. At 0.35 degrees of deviation, the receiver produces one dot deflection. At 0.7 degrees of deviation, the receiver produces a two-dot deflection.



ILS THEORY OF OPERATION



ILS TEST DISPLAYS

EHSI

Selecting an ILS mode on the EFIS control panel presents a navigation display with ILS information on the EHSI. This also provides an ILS mode discrete to the VOR control panel which causes paired tuning of the DME with the ILS frequency. The mode annunciation ILS appears on the lower left corner of the EHSI, and the ILS tuned frequency is displayed on the lower right.

Glide Slope Deviation

Glide slope deviation is displayed by a truncated triangle-shaped pointer moving against a four-dot scale. The glide slope index is a fixed small central rectangle which represents the airplane position relative to the glide slope beam (the pointer). One dot represents approximately 0.35 degrees deviation. Maximum deflection is +/- 2.2 dots.

Localizer Deviation

Localizer deviation is displayed by a bar symbol moving across a four-dot scale drawn as part of the selected runway course symbology located about the airplane present position symbol. One dot represents approximately one degree deviation. Maximum deflection is +/- 2.3 dots.

Selected Runway Course

The selected runway course pointer extends to the compass tape scale as shown. The course pointer is oriented to display the front course (FCRS) selected on the ILS control panel. The two course pointer segments also serve as a displacement index for the localizer deviation bar.

EADI

ILS glide slope and localizer deviation data are displayed on the EADI when a valid ILS frequency is tuned and the ground station transmission is received. Displays on the EADI are independent of the EHSI mode selector on the EFIS control panel.

Glide Slope Deviation

Glide slope deviation is displayed by a truncated triangle-shaped pointer symbol moving against a four-dot scale. The glide slope index is a fixed small central rectangle which represents the airplane position relative to the glide slope beam (the pointer). One dot equals approximately 0.35 degrees. Maximum deflection is +/- 2.2 dots (approximately 0.77 degrees). (The EADI glide slope display is the same as on the EHSI.)

Localizer Deviation

Localizer deviation is displayed by a rectangle-shaped pointer moving against a four-dot scale. The localizer index is a small fixed central rectangle which represents the airplane position relative to the localizer beam (the pointer). One dot equals approximately 1.0 degree. The maximum deflection is +/- 2.2 dots. Expanded scale - the standard four dot scale is replaced by the expanded two dot scale when LOC deviation data is used by the AFDS and the deviation is less than 0.625 degrees. For the expanded scale, one dot equals approximately 0.5 degree. The maximum deflection is +/- 1.25 dots.

Rising runway

At radio altitudes of 2500 ft or less, the rising runway is attached to the localizer pointer. At zero radio altitude, the rising runway touches the airplane symbol. At radio altitudes of 200 ft and greater, it assumes its lowest position.



ILS TEST DISPLAYS



ILS TEST SEQUENCE

Test Preparations

Apply power to the ILS, EFIS, and standby attitude systems. On the EFIS control panel, select ILS mode. On the standby attitude indicator, select ILS mode. On the ILS control panel, select any frequency other than park (frequency readout for park is ---.--).

On the ILS control panel, press and release the TEST push-button switch; or, on the receiver of the ILS system to be tested - left, center, or right - press and hold TEST push-button switch.

Display Sequence

The following test sequence will occur during the test and may be observed on the EADI, EHSI, and standby attitude/ILS indicator.

An invalid data display occurs for 3 seconds.

An NCD condition occurs for the next 2 seconds.

G/S and localizer pointers move to one dot up and one dot left respectively, for 3 seconds.

G/S and localizer pointers move to one dot down and one dot right for remainder of the test.

ILS Receiver Front Panel Test Sequence

The front panel LED’s will illuminate for three seconds, then extinguish for two seconds and then those will illuminate that indicate the status (as shown on the graphic). LED’s operate with ILS receiver test switch only.



ILS TEST SEQUENCE



MARKER BEACON INTRODUCTION

General

The purpose of the marker beacon system is to indicate to the flight crew that the airplane is passing over a particular geographical location (such as a point along an air route) or points along an instrument landing path. The ground stations transmits narrow beam RF signals modulated either 400, 1300, or 3000 Hz audio. As the aircraft flies over the a specific point, these signals turn on specific lights and audio.



MARKER BEACON INTRODUCTION



MARKER BEACON COMPONENTS

General

Marker Beacon Lights are on the P1 and P3 panels. While the Marker Beacon Receiver is a sub-section of the VOR Receivers (VOR/Marker Beacon Receivers) on the E2-2 and E2-3 racks. The antenna is located underneath the aircraft.

Operation

The antenna located on the underside of the fuselage, receives 75 MHz signals transmitted from a ground station and routes them to the marker beacon receiver located within the VOR/Marker receiver where the audio modulation is detected. The Marker Beacon Receiver module is only operational within the left VOR/Marker receiver. Flight deck indication providing visual identification of the beacon being flown over is provided by a set of marker beacon lights on each of the pilot's instrument panels. The blue light "OUTER" illuminates over the outer marker. The amber light "MIDDLE" illuminates when over a middle marker and the white light "INNER" illuminates when flying over either an inner marker or an airways marker. Aural identification is also provided by one of three audible tones sent to the audio selector panels. The outer marker is located approximately 4 miles from the runway end. When the airplane passes over this marker, the blue "OUTER" light on the P1 and P3 panels illuminates and a 400 Hz tone, keyed as continuous dashes, is heard on the flight interphone system. The middle marker illuminates the amber "MIDDLE" lights on the P1 and P3 panels, and a 1300 hz tone is keyed as alternate dots and dashes. This marker is located approximately 1/2 mile from the runway end.

The inner marker, located approximately .1 mile from the runway end, illuminates the white "INNER" lights on the P1 and P3 panels, and a 3000 hz tone is keyed as continuous dots. The back course marker is located at the opposite end of the runway from the inner marker at the typical final approach fix location. It also has a 3000 hz tone and illuminates the white "INNER" lights and 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.

Marker Beacon Receiver

The Active Marker Beacon is located in the left VOR/MKR Receiver. The receiver processes the signals and illuminates the appropriate light.

Antenna

The antenna is shaped like a canoe hull and is mounted on the bottom of the fuselage.

Marker Beacon Lights

The lights are colored blue, amber, and white and are labels OUTER, MIDDLE, and INNER.



MARKER BEACON COMPONENTS



MARKER BEACON TEST

Test

Press and hold TEST switch for at least 2-3 seconds on the VOR/MKR Receiver. All lights will illuminate and then extinguish. Then a status light will illuminate to indicate pass or failure.

• A green LRU STATUS LED indicates no faults. • A red LRU STATUS LED indicates VOR/MKR receiver failure. • The red CONTROL INPUT FAILED is relative only to VOR circuitry.

In flight compartment, the marker beacon lights illuminate simultaneously for approximately 10 seconds.



MARKER BEACON TEST



ATC INTRODUCTION

System description

The ATC system on the airplane consists of two transponders, two antennas and a dual ATC control panel. It provides altitude and identification reply signals to the interrogating ATC ground station. These signals are used to identify and locate an aircraft as it fly’s through each ATC ground sector. Only one transponder can be active and the other remains in the standby mode.



ATC INTRODUCTION



ATC GENERAL

General

The ATC System consists of:

• TOP and BOTTOM ATC Antennas • ATC Antenna switches • ATC/TCAS Control Panel • ATC Transponder



ATC GENERAL



ATC COMPONENT LOCATIONS

ATC Transponders

Located in the main equipment center, on rack E2-2 (left ATC) and E2-3 (right ATC).

Dual ATC Control Panel

Located on the aft pilot's control stand (P8).

ATC Antennas

Located on the forward fuselage at station 448 (top and bottom).

ATC Circuit Breakers

Located on the overhead circuit breaker panel P11; LEFT ATC, RIGHT ATC, ATC RF SW.

Antenna RF Switches

Located in the main equipment center, inboard on the E2 rack forward stanchion.

Air/Ground Relays

Located in the P36 panel (left ATC) and the P37 panel (right ATC).



ATC COMPONENT LOCATIONS



ATC FUNCTIONAL TEST

General Description

The dual mode S ATC/TCAS control panel provides control of the left and right mode S ATC transponders and the TCAS (traffic alert and collision avoidance system) receiver/transmitter. The TA and TA/RA positions of the rotary function switch apply to TCAS and will be covered in that book.

The control panel provides mode and transponder selection, identification code selection, and IDENT mode selection. Transponder failure is shown by the ATC FAIL light.

The control panel receives 115 volts ac from the left and right ATC transponder circuit breakers. Two separate power supplies provide power for two independent sets of electronics, one set for each ATC system.

Transponder and Mode Select

The rotary function switch enables the selected transponder in the selected mode of operation, places both transponders in the standby mode, or starts a test of the selected transponder. The standby mode inhibits transponder operation except for BITE functions. With the rotary function switch in the ALT ON, TA, or TA/RA positions, the transponder selected with the ATC 1/2 switch is made active with mode C altitude reporting enabled. With the function switch in the ALT OFF position altitude reporting is inhibited.

Identification Code and IDENT Mode Select

The four digits of the 4096 identity code are selected by the two sets of concentric rotary switches. The code (0000 to 7777 octal) is shown on an LCD display along with ATC 1 or ATC 2 to show which transponder is selected. An ATC IDENT push-button allows SPI pulses to be transmitted.

Test

Rotating the function switch to the spring loaded TEST position will start a BITE test of the selected transponder. The ATC FAIL light (and EICAS ATC FAIL message) will show momentarily after an all-segment display is shown in the display window.

The ATC functional test provides a quick check of the system using the transponders BITE capabilities, and does not require extra test equipment.

Test Preparation

Power must be applied to the ATC and ADIRU.

Test Starting

Test starting at both the transponders' front panel in the main equipment center and from the transponder control panel on the flight deck are shown on the graphic.

Test Indications and Results

The normal and abnormal indications on the transponder front panel are shown on the graphics. See ATC transponder for a complete explanation of status lamp meaning.



ATC FUNCTIONAL TEST



TCAS SYSTEM INTRODUCTION

TCAS System Components

The TCAS System is composed of a TCAS Processor, top and bottom mounted TCAS antennas, two Mode S Transponders, a combined Transponder/TCAS Control Panel, two Traffic Alert/Vertical Speed Indicators (TA/VSI), and aural warning.

Using the directional antenna, the TCAS Processor interrogates other aircraft transponders and performs calculations necessary to identify potential conflicts. When potential conflicts exists, the processor provides an aural alert and activates TCAS displays.

During TCAS advisories, the inside of the VSI becomes a traffic display. IF an actual conflict develops, colored arcs are displayed on the TA/VSI and another aural alert is provided. These arcs identify the vertical speed required to insure separation and the aural alert reinforces the required action.

A combined control panel is used to operate both the Mode S Transponder and the TCAS system. In addition to the traditional ATC Mode C function, the Mode S Transponder will communicate and coordinate avoidance maneuvers if conflicting traffic is TCAS equipped.

In addition, weather radar indicators provide full time display of traffic when selected.

TCAS Aircraft Interface

The TCAS system is interfaced with the following aircraft systems:

• Radio Altimeter. • Air Data Computer. • Inertial Reference System. • Air/Ground Sensor. • Gear Position. • GPWS. • Windshear System. • Aural Warning System.

GPWS or windshear commands will inhibit TCAS aural alerts and cause the TCAS system to revert to the TA only mode.

Aircraft performance capability is not directly interfaced to TCAS. In providing vertical avoidance advisories, TCAS does not know existing performance margins, or aircraft performance degradation's (i.e., engine out).



TCAS SYSTEM INTRODUCTION



TCAS DISPLAY SYMBOLOGY AND INDICATIONS

Symbology

• HOLLOW WHITE DIAMONDNon-threatening traffic

• SOLID WHITE DIAMONDNon-threatening traffic (This symbol is filled in because the aircraft is in close proximity. This is defined as within 1200 feet and six (6) nm.)

• SOLID YELLOW CIRCLEIdentifies traffic causing a TRAFFIC ALERT (potential conflict).

• SOLID RED SQUAREIdentifies traffic which is causing a RESOLUTION ADVISORY. (Immediate threat)

Data Tags

Altitude of displayed traffic is shown as the difference between your altitude and that of the traffic. This relative altitude is represented as a two digit number indicating hundreds of feet, (i.e., 05 = 500 feet).

A plus or minus sign and the placement of the altitude information are both used to indicate whether displayed traffic is above or below.

• -08 - is a traffic advisory eight hundred feet below your altitude. • +08 - is a traffic advisory eight hundred feet above your altitude.

Symbols displayed without altitude information indicate no altitude received. Yellow circles with no altitude information may be a significant hazard, yet TCAS is unable to provide Avoidance Advisories without altitude information.

Aircraft that are climbing or descending in excess of 500 feet per minute (FPM) will be displayed with a data tag which includes an arrow pointing in the appropriate direction.

-15 - is a traffic advisory 1500 feet below and climbing at a rate that exceeds 500 FPM.

EHSI

The two Electronic Horizontal Situation Indicators (EHSI’s) are the primary indicators for TCAS display. They function as normal EHSI’s until traffic is detected at which time the center of the instrument "pops up" to display traffic. The display is fixed in range, showing traffic approximately six miles in front of an aircraft like symbol representing present position. Ranges to the side and behind are to scale. There is a two mile range ring that surrounds our aircraft symbol.

Red and green colored arcs are displayed around the outside of the EHSI. These arcs identify the vertical speed required to insure proper separation. Both colors and shapes are used to show other aircraft.

Traffic Display on EHSI

TCAS traffic will be displayed when the "TCAS" or "TCAS/WX" mode is selected on the color weather radar. Unique TCAS functions on the EHSI are the A/B button and the FL button. The A/B button is a push-button which will bias the altitude band displayed on the indicator. Possible selections are:

ABOVE 8700 feet above to 2700 below.NORMAL 2700 feet above to 2700 belowBELOW 2700 feet above to 8700 below

WX/TCAS mode on EHSI

All ranges may be selected. TCAS may display traffic at distances up to 40 nm. At higher ranges TCAS information will be compacted in the lower display area. The five (5) mile range, while clearly displaying TCAS traffic, will not display weather. In this display mode conflicting aircraft that approach from behind can not be displayed. An annunciation will indicate "intruder behind".



TCAS DISPLAY SYMBOLOGY AND INDICATIONS



TCAS OPERATION SURVEILLANCE AREA

General

Maneuver indications on the electronic attitude direction indicators (EADIs) guide the flight crew to avoid a possible collision.

Aurals come from the aural warning speakers.

System Inhibits

TA’s or RA’s may not occur if the airplane radio altitude is too low or if a higher

priority alert exists.

This table shows the inhibits and the parameters that cause the inhibits:

CONDITION PARAMETER

Increased

descent RA Inhibited below 1450 ft AGL

Descend RA

Inhibited below 1000 ft AGL in descent and

1200 ft AGL in climbResolution

Advisories

Inhibited below 900 ft AGL in descent and below

1100 ft AGL in climb. (TCAS automatically goes into

TA ONLY)TA voice

message

Inhibited below 900 ft AGL in descent and below

1100 ft. AGL in climbGPWC alerts Inhibits RAs

PWS alerts Inhibits RAs



TCAS OPERATION SURVEILLANCE AREA



TCAS OPERATIONAL TEST

General

The TCAS and ATC systems can be tested from the flight station or from the face of their respective computers.

EADI Resolution Advisory

On the EADI a Resolution advisory shows as red brackets pointing up for an RA UP ADVISORY for the pilots, and red brackets pointing down for a RA DOWN ADVISORY.



TCAS OPERATIONAL TEST



TCAS ANTENNA LOCATIONS

General

The Top TCAS antenna is located at station 380 and the Bottom TCAS antenna is located at station 399. Both antennas are directional until the landing gear discrete is received from the landing gear module. At this time the lower TCAS antenna becomes an omnidirectional antenna to prevent blind spots.



TCAS ANTENNA LOCATIONS



TCAS TEST DISPLAYS

General

The following illustration shows EHSI and EADI displays during TCAS Test.



TCAS TEST DISPLAYS



WEATHER RADAR INTRODUCTION

General

The weather radar system provides the pilots with visual indication of storm conditions in order to avoid heavy precipitation or turbulence.

As a secondary function, the weather radar antenna can be tilted downward to provide a display of significant land contours. This can be useful as an additional navigation aid. The radar system generates RF pulses and transmits them through the antenna. Pulses reflected from targets are received by the antenna and processed by the system for display. The receiver processes the return signals and sends them to the EFIS symbol generators for display on the EHSI and to the weather radar indicator.

Controls for the system are on the EFIS control panel (for weather radar enable/disable for EHSI display and range), and on the weather radar indicator (for on/off, mode, range, gain, and tilt). The antenna is stabilized using the airplane attitude (pitch and roll from IRS) to ensure that the scan is parallel to the horizon.



WEATHER RADAR INTRODUCTION



WEATHER RADAR - OVERVIEW

System Components

The weather radar system consists of a one X-band transceiver located in the forward equipment bay, a flat plate weather radar antenna located under the nose radome. The system also interfaces with the EFIS system for weather radar displays on the EHSI’S weather images are able to be displayed on the EHSI. The EFIS control panel provides enable/disable and range data for display of weather data on the respective EHSI in MAP, EXP MAP, or EXP ILS modes only.

System Outputs

Display data from the transceiver is sent to the EFIS symbol generators and to the indicator. In addition to the weather data, the transceiver sends system status messages, including faults. Three data buses are necessary since for each display unit an individual range can be selected. The EFIS symbol generators format the data from the transceiver for presentation on the EHSI's.

The EFIS source select switch determines which symbol generator supplies the EHSI's. The indicator formats the data for presentation on its own. In the "WX" (weather avoidance) mode the weather targets are color-coded by the intensity of the return. The display correlation to approximate rainfall is as follows:

• Black - very light or no returns = less than 0.7 mm/hr. • Green - light returns = 0.7 - 4 mm/hr. • Yellow - medium returns = 4 - 12 mm/hr. • Red - strong returns = greater than 12 mm/hr. • Magenta - very strong returns = greater than 25 mm/hr.

In the "MAP" mode the weather radar system can be used in ground mapping to identify terrain features. The display colors in the "MAP" mode are the same as in the "WX" mode.

To operate in the "TURB" (turbulence detection) mode the weather radar system requires the presence of precipitation, therefore, turbulence detection does not display clear air turbulence. Turbulence information is limited to the first 40 nautical miles. Turbulence within this range will be displayed in magenta along with the weather displayed in red, yellow, and green, on the weather indicator only. Only weather will be displayed beyond the 40 nm turbulence limit when a range of more than 40 nm is selected.



WEATHER RADAR OVERVIEW



WEATHER RADAR SYSTEM COMPONENTS

Weather Radar Transceivers The weather radar transceivers transmit and receive rf pulses. The returning pulses are converted into video data and sent to the EFIS symbol generators for display on the EHSI, and to the weather radar indicator.

The transceivers also compute stabilization corrections and sends them to the antenna to maintain a horizontal scan regardless of the airplane roll or pitch angle.

The transceiver also monitors the entire system. Status and fault words are sent along with the weather data to the EFIS symbol generators. The fault word isolates the fault to the LRU.

The "TEST" switch places the system into a test mode in the same manner as the "TEST" position of the mode select switch on the weather radar control panel. In addition, the "TEST" switch on the transceiver momentarily turns on all led indicators to verify all are working, then extinguishes all LED's except those displaying existing faults.

There are two tranceivers located in the pedestal or base of the antenna. Unlike previous systems that were located in the fwd equipment compartment.

Weather Radar Processors

The two weather radar processors are located in the forward equipment compartment. These processors take the data coming from the weather radar tranceivers and compute the information for display on the L & R EHSI’s or LDS displays..

Antenna

The antenna is used to radiate a beam of energy and then receive it back. The antenna scans 90 degrees parallel to the horizon at a rate of 15 looks per minute. The stabilization is automatically adjusted as the aircraft changes attitude. The tilt no longer needs adjusted due to the auto scan of weather in front of the aircraft from 0 to 60000 feet. This paints a more accurate picture of upcoming weather patterns.

L & R Electronic Horizontal Situation Indicators

The weather radar control panel located on the left side of the center pedestal, supplies the transceiver with mode control, selected tilt angle, range control, receiver auto/manual gain control, and indicator brightness control.

The mode select switch turns on the transceiver when any mode is selected except "OFF".

The range selector selects the range for the data to be displayed on the indicator. The range may be different from those selected on the EFIS control panel for display on a EHSI.

The "AUTO" position on the gain control provides a preset, calibrated receiver gain level. The "MIN" position indicates the minimum gain for manual control.

The "WX ON" switch on each EFIS control panel enables/disables weather radar to be displayed on the on-side EHSI, providing the weather radar system has been turned on by the mode select switch on the weather radar indicator. The selected range from each EFIS control panel tells the transceiver how to process the data received from returning pulses to be displayed on the respective EHSI. Each pilot can select a different range.



WEATHER RADAR SYSTEM COMPONENTS

WAVEGUIDE RUNS REPLACEDBY COAXIAL CABLE ON THERDR‐4000 SYSTEM

WX RADAR XCVRSARE NOW PART OF THE PEDESTAL

AND RADAR PROCESSORIS NOW IN

FWD EQUIPMENT CENTER



WEATHER RADAR DISPLAYS

General

Weather radar system status and fault annunciation’s are displayed on the EHSI on two lines in the lower left corner of the display. The displayed messages and their meaning are listed in a chart on the graphic.

The message "WXR DSPY" indicates an EHSI overheat condition. 30 seconds

after the temperature reaches 75oc, the raster, and consequently, the weather display, is removed. When the condition is corrected, they return. If, on the

contrary, the temperature reaches 100oc, the entire display is removed. The message "DSPY" has priority over all other messages.

Range Disagreement Annunciation

When the symbol generator senses disagreement between the EFIS control panel range and the transceiver range, the annunciation "wxr range disagree" is annunciated. This annunciation is possible in the following EFIS modes: "EXP VOR", "EXP ILS", or "MAP" modes.

When both the transceiver and the FMC ranges disagree with the control panel range the message is "WXR/MAP RANGE DISAGREE". This annunciation appears only when the mode selected on the EFIS control panel is "map".

Weather Radar Indicator Fault Annunciations

Status messages are displayed with the weather display not being removed. The messages and their meanings are listed on a chart on the graphic. The message "STAB" and "CAL" each have two meanings, depending on the color of the message display.

Fault Messages Fault messages are displayed with the weather display being removed. The messages and their meanings are also listed on the graphic. A detected attitude, indicator, calibration, or cooling fault is annunciated only in the "TEST" mode.

Test Initiation

The flight deck self-test is initiated by positioning the mode selector switch on the weather radar indicator to "test". The test results are annunciated on the weather radar indicator and on the EHSI's whenever their respective EFIS control panels' "WXR" switch is in the "ON" position and the "EXP VOR", "EXP ILS", or "MAP" mode is selected.

Weather Radar Indicator Test

Test pass display:

A red/yellow/green rainbow shaped test pattern with a magenta colored wedge in the center appears and the message "TEST" is displayed in the upper left corner. If the turbulence mode is not active in the transceiver, the magenta wedge will not appear in the center of the test pattern.

Test fail display:

The test pattern is blanked and the LRU’s responsible for the faults appear on the center of the display. The meanings of the individual messages is shown on a chart on the graphic.



WEATHER RADAR DISPLAYS



REACTIVE AND PREDICTIVE WINDSHEAR

Windshear Alerts

The weather radar transceiver and ground proximity warning computer generate windshear alerts.

The weather radar transceiver makes a windshear alert within three NM of a windshear. This is a predictive windshear. Between 3 NM and 1.5 NM, a caution is made. Between 1.5 NM up to the windshear, a warning is made.

The ground proximity warning computer makes windshear warnings when in the windshear. This is called a reactive windshear.

PWS Warning - Flight Deck Effects

A PWS warning has these visual and aural annunciations:

• A red WINDSHEAR message on the EADI • A red WINDSHEAR message on the EHSI • A windshear symbol on the EHSI • Master warning lights on • Aural annunciation - “WINDSHEAR AHEAD” (takeoff) • Aural annunciation - “GO AROUND, WINDSHEAR AHEAD” (approach)



REACTIVE AND PREDICTIVE WINDSHEAR



PREDICTIVE WINDSHEAR ALERTS AND DISPLAYS

Windshear Alerts

The weather radar transceiver and ground proximity warning computer generate windshear alerts.

The weather radar transceiver makes a windshear alert within three NM of a windshear. This is a predictive windshear. Between 3 NM and 1.5 NM, a caution is made. Between 1.5 NM up to the windshear, a warning is made.

The ground proximity warning computer makes windshear warnings when in the windshear. This is called a reactive windshear.

PWS Warning - Flight Deck Effects

A PWS warning has these visual and aural annunciations:

• A red WINDSHEAR message on the EADI • A red WINDSHEAR message on the EHSI • A windshear symbol on the EHSI • Master warning lights on • Aural annunciation - “WINDSHEAR AHEAD” (takeoff) • Aural annunciation - “GO AROUND, WINDSHEAR AHEAD” (approach)

Alert Prioritization

PWS caution and warning alert are prioritized with other flight deck caution and warning level conditions.

To prevent conflicts or simultaneous voice and visual alerts to the flight crew, the GPWS prioritizes the alert messages.

If the GPWS finds that the PWS warning overrides the current GPWS alert, the inhibit discrete is removed.

The PWS audio alert inhibit is also used to mix PWS alerts with other higher priority alerts that have time to allow alerts to be annunciated.

PWS Caution - Flight Deck Effects

A PWS caution has these visual and aural annunciations:

• Amber WINDSHEAR message on the EHSI • Windshear symbol on the EHSI • Aural annunciation - “MONITOR RADAR DISPLAY”



PREDICTIVE WINDSHEAR ALERTS AND DISPLAYS



WEATHER RADAR TEST DISPLAYS

Test Initiation

The flight deck self-test is initiated by positioning the mode selector switch on the weather radar indicator to "test". The test results are annunciated on the weather radar indicator and on the EHSI's whenever their respective EFIS control panels' "WXR" switch is in the "ON" position and the "EXP VOR", "EXP ILS", or "MAP" mode is selected.

Weather Radar Indicator Test

Test pass display:

A red/yellow/green rainbow shaped test pattern with a magenta colored wedge in the center appears and the message "TEST" is displayed in the upper left corner. If the turbulence mode is not active in the transceiver, the magenta wedge will not appear in the center of the test pattern.

Test fail display:

The test pattern is blanked and the LRU’s responsible for the faults appear on the center of the display. The meanings of the individual messages is shown on a chart on the graphic.



WEATHER RADAR TEST DISPLAYS



PREDICTIVE WINDSHEAR INTERFACE

General Description

The Predictive Windshear (PWS) part of the Weather Radar (WXR) system interfaces with other airplane systems and components.

28vdc WXR RT Circuit Breaker

If the weather radar transceiver loses power, the 28vdc circuit breaker enables the PWS FAIL discrete to still be sent.

Proximity Switch Electronic Unit

The proximity switch electronic unit sends an analog discrete to PWS for landing gear lever position. PWS uses this discrete in its takeoff/approach alert logic.

EICAS Computers

Take-off engine thrust from the left and right EICAS computers goes to the PWS as the A qualifier for the PWS. This allows radar operation on the ground even if not selected on from the EFIS control panel.

A discrete from the weather radar transceiver goes to the EICAS Computers when PWS fails.

Radio Altimeter

The Radio Altimeter (RA) provides radio altitude data to PWS. PWS uses this data for these functions:

• Turn PWS on and off • Enable/disable display and alert functions

Air Data System

The Air Data Computer sends airspeed data to PWS.

Aural Warning Speakers

Audio warnings go directly to the aural warning speakers.

Ground Proximity Warning Computer

PWS sends windshear alert data to the Ground Proximity Warning Computer (GPWC) on the ARINC 429 hazard bus. The GPWC prioritizes alerts. If a higher priority alert exists, the GPWC sends an inhibit discrete to WXR. The discrete inhibits PWS aural alerts if the GPWS alerts are a higher priority.The PWS aural alert stops when a higher GPWS alert is received.

Symbol Generators

A discrete goes from the weather radar transceiver goes to the Symbol Generators to make a display for a PWS caution or alert.

A discrete goes from the weather radar transceiver goes to the Symbol Generators when PWS fails.

Traffic Alert and Collision Avoidance System

WXR sends an inhibit discrete to the Traffic Alert and Collision Avoidance System (TCAS) computer when there is a PWS alert. This discrete does these functions:

• Changes Resolution Advisories (RAs) to Traffic Advisories (TAs)Inhibits all TCAS audio alerts



PREDICTIVE WINDSHEAR INTERFACE

RADAR PROCESSOR



EGPWS INTRODUCTION

General

The ground proximity warning computer (GPWC) establishes the limits for the ground proximity mode and windshear envelopes. The computer compares the flight path and terrain clearance status to the mode limits for the airplaneconfiguration to find if there is an alert or warning condition.



EGPWS INTRODUCTION



ENHANCED GROUND PROXIMITY WARNING SYSTEM INTRO-DUCTION

General

The enhanced ground proximity warning system computes two levels of terrain alerting envelopes: caution and warning.

Terrain display colors indicate the height of the terrain relative to the current airplane altitude.

The enhanced ground proximity warning system, look-ahead caution alert includes unique voice aural, terrain display, and map annunciations.

The enhanced ground proximity warning system look-ahead warning alert aurals are nearly identical to those for the basic GPWS warnings in order to elicit the same pilot response. In addition, threatening terrain is presented on the map display.



ENHANCED GROUND PROXIMITY WARNING SYSTEM INTRODUCTION



EGPWS - GENERAL DESCRIPTION

General

The GPWS has different modes of operation to detect unsafe conditions in flight. These are the modes and the conditions that cause them to operate:

• Terrain mode - insufficient terrain clearance from GPWS stored terrain • Terrain clearance - insufficient terrain clearance during approach • Mode 1 - excessive descent rate • Mode 2 - excessive closure rate with terrain • Mode 3 - excessive altitude loss during climb-out • Mode 4 - insufficient terrain clearance • Mode 5 - excessive deviation below glideslope • Mode 6 - aural callouts • Mode 7 - windshear conditions

The GPWC makes synthesized voice messages and sends them through the warning electronics unit to the aural warning speakers.

These annunciations alert the crew to active GPWS modes and hazardous conditions:

• Master warning lights, red PULL UP light, and red WINDSHEAR light • WINDSHEAR annunciation on the EADI • GND PROX/G/S inhibit switch light

The GPWS interfaces with these systems to determine warnings and cautions:

• Global positioning system • Weather radar system (WXR) • EFIS control panel • Radio altimeters • Air data inertial reference unit (ADIRU) • Flight management computer (FMC) • Stall warning system • Landing gear handle • Flap/stab position module • ILS receiver

The GPWC receives a discrete when the flight crew selects TERR on the EFIS control panel. The GPWC controls the terrain relays and sends the terraindisplay to the EFIS symbol generators. The GPWC can also turn on the terrain display automatically.

EFIS Control Panel Terrain/Weather Switches

Both terrain and weather are selectable thru switches located on the EFIS Control Panels. However both modes cannot be displayed at the same time. If both are selected then Terrain has priority and will be displayed. However, if Predictive Windshear were to detect a microburst with respect to the flight path of the aircraft then it will override the Terrain function and display the windshear icon on the EHSI.



EGPWS - GENERAL DESCRIPTION



ENHANCED GROUND PROXIMITY WARNING COMPUTER

Purpose

The GPWC contains a worldwide terrain database and an airport database. Both databases store data by latitude and longitude. The GPWC uses GPS position and, if the GPS fails or is NCD, the GPWC uses inertial position.The GPWC compares the airplane position, track, and altitude to the co-ordinates in the database to determine if an alert condition exists.

The GPWC also has a database that contains the location of all hard surface runways in the world that are longer than 3,500 feet. The GPWC compares airplane latitude, longitude, and radio altitude with an envelope around the approach runway. If the airplane descends through the floor of the envelope, the GPWC gives an alert.

The GPWC monitors the weather radar hazard bus. If the WXR has a higher priority message than the GPWS, the GPWC makes sure the weather radar message shows. If the GPWS has a higher priority message than the weatherradar, the GPWC makes sure the GPWS message shows. The GPWC also sends a WXR predictive windshear inhibit when the GPWS has a higher priority message.

Terrain or weather data will be displayed on the EHSI.

Physical Description

The GPWC is a 2 MCU chassis and weighs seven pounds (3.2 kg).

Front Panel

The GPWC front panel has three status LEDs and a door. These are the status LEDs on the front panel:

• EXTERNAL FAULT - amber LED • COMPUTER OK - green LED • COMPUTER FAIL - red LED.

The front panel door allows access to these:

• PRESS TO SELF-TEST switch - starts a GPWS test • Headphone jack - lets you hear self-test audio • Memory card slot - lets you upload software from a Personal Computer • Memory Card International Association (PCMCIA) memory card or

download fault and warning history data • Upload/download status indicators - shows conditions of upload or

download operation • RS-232 connector - used for shop test or for the upload/download of data



ENHANCED GROUND PROXIMITY WARNING COMPUTER



EGPWS - COMPONENT LOCATIONS

GND PROX - G/S INHB Light Switch

The GND PROX-G/S INHB light/switch has a dual function. The amber light shows alerts caused by modes 1 through 5. The switch function prevents or cancels mode 5 annunciations.

If you push the switch before the mode 5 indications start, the annunciations will be inhibited. If you push the switch after the indications start, the annunciations will be cancelled.

Flap and Gear Override Switches

The ground proximity flap override (GND PROX FLAP OVRD) switch and the ground proximity/configuration gear override (GND PROX/CONFIG GEAR OVRD) switch let the crew simulate flaps down 25 units or more or landinggear down positions. These are guarded alternate-action push button switches. When the override function is on, the switches are white. Push the switch again to cancel the override function.

Terrain Switch

The terrain switch on the EFIS control panel lets the flight crew enable the automatic terrain display and terrain alerts features on the onside ND.

Terrain Override Switch

The terrain override switch lets the flight crew inhibit the automatic terrain display feature on the ND.

Ground Proximity Test Switch

The ground proximity test switch starts the GPWS self- tests. When you use the switch, the confidence test starts. If you hold the switch through the windshear annunciations, the full vocabulary test follows the confidence test.

Terrain Awareness Display Data

The EHSIs show this GPWS data:

• Terrain awareness display • GPWS system messages • GPWS alert messages.

If there is a terrain caution alert, the terrain awareness display changes from dots to a solid yellow. If there is a terrain warning alert, the terrain awareness display changes from dots to a solid red.

GPWS System Messages

These GPWS system messages show in cyan on the EHSIs:

• TERR shows when terrain data shows • TERR TEST shows when the GPWS is in the self-test mode

GPWS Alert Messages

These GPWS alert messages show on the EHSIs:

• TERRAIN (red) shows a terrain warning • TERRAIN (amber) shows a terrain caution



EGPWS - COMPONENT LOCATIONS



GPWS LIGHTS AND SWITCHES

Purpose

The purpose of GPWS annunciator lights is to provide a visual indication of GPWS modes, and that of the flap and gear override switches to provide the capability of simulating a flap-down and/or gear-down condition.

WINDSHEAR Light

The red WINDSHEAR light is located on the captain's instrument panel P1-3.

PULL UP Light

The red PULL UP light indicates a mode 1 or mode 2 pull up warning condition. It is located on the captain's instruments panel P1-3.

GND PROX - G/S INHB Light Switch

The GND PROX-G/S INHB light/switch has a dual function. Its amber light is used to annunciate alerting modes 1 through 5. Its switch inhibits or cancels mode 5 (below glide slope) when actuated below 1000 feet on approach. If the switch is pressed before the mode 5 indications have started, the indications - visual and aural - will be inhibited. If the switch is pressed after the indications have started, the indications - visual and aural - will be canceled. Once canceled or inhibited, the indications cannot be reinstated or rearmed simply by a repeated switch actuation. Mode 5 is automatically rearmed when the airplane descends below 30 feet or climbs above 1000 feet radio altitude.

The switch is a momentary switch and is located on the captain's instrument panel P1-3.

Flap and Gear Override Switches

The GND PROX FLAP OVRD (ground proximity flap override) switch and the GND PROX/ CONFIG GEAR OVRD (ground proximity/configuration gear override) switch serve to simulate flaps down 25 units or more or landing gear down positions, respectively. These are guarded alternate-action push-button switches. When the override function has been activated, the switches illuminate white. A repeated switch actuation cancels the override function. Both override switches are located on the first officer's instrument panel, P3-1.

The effect of override switch actuation will be described later.

Ground Proximity Test Switch

The ground proximity test switch is used to initiate ground proximity flight compartment self tests. When the switch is actuated, the confidence test is initiated. If the switch is held through the windshear annunciation, the full vocabulary test will follow the confidence test.

The switch is located on P61.



GPWS LIGHTS AND SWITCHES



GPWS - FUNCTIONAL DIAGRAM

Interfaces

Input Signals - The interface systems shown provide data, status, and discrete inputs to the ground proximity warning computer. The specific types of signals from each data source are summarized in GPWC - COMPUTATION DATA SOURCES.

SIGNAL DESTINATION

Aural messages Left and right aural warning siren/owl module (warning electronics unit).

Windshear Master warning module (warning discrete electronics unit); WINDSHEAR light; left, center and right EFIS symbol generators and TCAS computer.

SIGNAL DESTINATION

Mode 1 and 2 PULL UP light and PULL UP warning master warning discrete module (warning electronics unit) and TCAS computer.

Mode 1 through 5 GND PROX-G/S INHB alerting light and TCAS discretes computer.

Ground proximity digital flight mode (serial acquisition unit. message).

Warning Electronics Unit Outputs - The aural messages are amplified in the left and right siren/owl modules then routed to the captain's and first officer's aural warning speakers. On command from the GPWC, the master warning module turns on the captain's and first officer's master warning lights.

Operation

GPWC Functions - The computer uses data inputs to compute the airplane flight status in relation to mode 1, 2, 3, 4, and 5 mode boundaries. When required, the computer generates the annunciation signals, and transmits these signals to the visual and aural annunciation devices.

Order of Priority of Messages - If more than one ground proximity warning system mode occurs at the same time, only the one having the highest priority is annunciated. The order of priority is as follows:

• "WINDSHEAR" • "WHOOP WHOOP PULL UP" • "TERRAIN, TERRAIN" • "TOO LOW • “TERRAIN" • "TOO LOW - GEAR" • "TOO LOW • “FLAPS" • "SINK RATE" • "DON'T SINK" • ”GLLIDE SLOPE" • RA CALLOUTS

Program Pins - Program pins are jumpered to the program pin common to provide functions (such as airplane and aural vocabulary selection).

Fault Monitoring:

• Fault Storage - BITE circuitry performs continuous and periodic checks of internal circuits and input data. Detected faults are stored for the last 10 flights in a nonvolatile fault memory for later readout. The memory can only be cleared during bench test.

• Fault Display - When actuating the STATUS/HISTORY switch on the GPWC front panel, present faults or faults stored in fault memory are displayed in the BITE display window.

• EICAS BITE Message - The GPWS EICAS message "GND PROX BITE" indicates a fault condition in the system. It is a status and maintenance message and is displayed on the lower EICAS display unit. Presence of a new status message is indicated by the "STATUS" cue on the lower EICAS display unit if page is displayed.



GPWS - FUNCTIONAL DIAGRAM



EGPWS MODES

General

Mode 1

The mode 1 sink rate detector compares the actual radio altitude of the airplane with the threshold values of the mode 1 envelope as defined for the measured airplane barometric descent rate of the airplane.

Mode 2a

Closure rate detector functions - the closure rate detector first computes the instantaneous terrain closure rate of the airplane. It then compares the actual radio altitude of the airplane with the threshold values of mode 2a.

Mode 2b

If the flaps are down 25o or more, mode 2b applies. To compute mode 2b and to generate the respective signals for mode 2b annunciation, the closure rate detector uses the same method as in mode 2a.

Mode 3

When the airplane starts losing baro altitude, the altitude loss detector senses negative barometric descent rate and stores in a memory the altitude at which the descent started. When the altitude loss exceeds the threshold value defined for the given altitude, mode 3 is annunciated.

Mode 4

The altitude loss detector continuously transmits to the mode 4 (terrain closure) detector the calculated mode 3 critical alerting threshold value. When this threshold value reaches the upper boundary of the mode 4a envelope, mode 4 becomes armed and mode 3 disabled. The terrain closure detector compares the actual radio altitude with the boundaries of the mode 4 envelopes which depend on airspeed, flap and landing gear positions. If less than the terrain clearance defined by the envelope, signals for mode 4 annunciation are generated.

Mode 5

The glide slope detector verifies landing gear down and compares any measured deviation below the glide path with the threshold values as defined by the mode 5 envelope for the low-level audio and the normal-level audio as a function of radio altitude. Magnetic heading of the airplane is compared with

selected runway heading; if the difference is larger than 90o (back course), mode 5 is inhibited.

Mode 6

When the decision height alert detector senses transition through the selected decision height value (decision height alert discrete - ground), and if the radio altitude is less than 1000 feet and more than 50 feet and the gear is down, it generates a discrete for mode 6 annunciation.

Mode 7

During takeoff or landing configuration a vertical or horizontal windshear component detects a WINDSHEAR condition, then signals for mode 7 are generated.

Override Capabilities

• Grd Prox - G/S Inhb Light Switch

The GND PROX-G/S INHB light/switch has a dual function. It's amber light is used to annunciate alerting modes 1 through 5. Its switch inhibits or can-cels mode 5 (below glide slope) when actuated below 1000 feet on approach.

• Flap and Gear Override Switches

The Gnd Prox Flap Ovrd Switch and the Gnd Prox/Config Gear Ovrd Switch serve to simulate flaps down 25 units or more or landing gear down positions, respectively.



EGPWS MODES



EGPWS MODE ENVELOPE MODULATION

General

The purpose of envelope modulation is to modify specific warning and alerting envelopes to prevent nuisance GPWS mode annunciation’s in localities with marginal ground proximity terrain conditions in approach or take-off. The operation of envelope modulation is illustrated on the graphic by means of an example; the backcourse approach (runway 34) to Reno, Nevada.

Method Used In Envelope Modulation

40 airports throughout the world have been identified as having approach or departure peculiarities that are likely to produce nuisance annunciation. Reno, Nevada; San Diego, California; Ontario, California; Seoul, Korea; and Taipei, Taiwan. The critical areas have been defined by means of latitude and longitude data stored in the GPWS memory. When the GPWC senses the airplane's approach to such an area (magnetic track), 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 envelope of one or more GPWS modes. If the conditions are "right", the GPWS concludes that the "key fits" and, consequently, proceeds with the predetermined modulation. If the "key" does not "fit", no envelope modulation takes place.

When no glide slope signal is available - as in the case of a backcourse approach - a "snapshot" check is made. This check consists of determining the elevation of a "snapshot area" by subtracting its radio altitude from its corrected baro altitude and, consequently, by 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 its latitude/longitude coordinates and is situated a short distance before the envelope modulation area.

Envelope modulated airports

Alienate, Spain -mode 2AAgana Nas, Guam isl - mode 2AAlice Springs, Australia mode 2ACairns, Australia - mode 2A, 4Canberra, Australia - mode 2ACoolangatta, Australia - mode 2A, 4Cuenca, Ecuador - mode 2AGeneva, Switzerland - mode 2AHiroshima, Japan - mode 2AHobart, Tasmania - mode 2A, 4Hong Kong, B.C.C. - mode 1, 2AHot Springs, Virginia - mode 5, 6Kagoshima, Japan - mode 5, 6Launceston, Tasmania - mode 2ALeeds/Bradford, U.K. - mode 2A, 2BLisbon, Portugal - mode 2 2ALuxembourg, Luxembourg - mode 2A, 4Malaga, Spain - mode 2AMelbourne, Australia - mode 2ANice, France - mode 2A, 4Nome, Alaska - mode 2ANorth Bay, - Ontario - mode 2ANurnburg, Germany - mode 2AOntario, California - mode 2APaine Field, Washington - mode 5, 6Quito, Ecuador - mode 2A, 2B, 4Reno, Nevada - mode 2A, 4San Diego, California - mode 1Seoul, Korea - mode 2A, 4St. John's, Newfoundland - mode 2AStephenville, Newfoundland - mode 1Taipai (Sungshan), Taiwan - mode 2AUnalakleet, Alaska - mode 2AVagar, Faroe Islands - mode 2A, 4Victoria, B.C. - mode 2AWellington, New Zealand - mode 2AWrangell, Alaska - mode 2AZurich, Switzerland - mode 2A, 2B



EGPWS MODE ENVELOPE MODULATION



WINDSHEAR MODE

General

Windshear detection is a system incorporated within the ground proximity warning system computer (GPWC).

The windshear system is composed of three elements: Detection, Alert, and Guidance. A windshear condition is detected by comparing total aircraft energy with horizontal and vertical wind energy. Wind and inertial airplane information from the inertial reference systems (IRS), pitot/static information from the air data computers (ADC) is used by the ground proximity warning computer (GPWC) to determine a windshear condition.

If a windshear is detected, the ground proximity warning computer will initiate the warnings. The aural alert consists of an aural two-tone attention-getting sound (siren) immediately followed by a voice annunciation of "windshear, windshear, windshear". The aural alert only sounds once.

Visual cues are provided by the illumination of the master warning lights and a red "windshear" annunciator light on the P1-3 panel. The "WINDSHEAR" also appears in red at the bottom of the EADI's and remains in view until the windshear condition is no longer present. Only windshear conditions that approach the limiting performance capabilities of the airplane will initiate the warnings. The ground proximity windshear warning is only armed to activate from ground level to 1500 feet radio altitude. On take-off, the system arms climbing through 50 feet radio altitude.

The windshear guidance systems begins by displaying, whenever the flaps are not retracted, a Pitch Limit Indicator (PLI) on each EADI. The PLI, which comes to the EFIS symbol generators from the stall warning computer modules via the GPWC, indicates the pitch attitude at which stick shaker operation will occur for the existing flight conditions (AOA, Airspeed, Flap position). The distance between the PLI and the airplane symbol on the EADI represents the pitch margin between the current flight conditions and the stick shaker activation point.

When encountering a windshear during take-off or approach, severe enough to activate the windshear warnings, pushing a "go-around" switch engages the autothrottle and autopilot/flight director systems in a windshear recovery guidance mode.

The flight director pitch command bar will smoothly transition from a speed

mode to an attitude mode and command a pitch attitude of 15o up, or

approximately 1o below the PLI, whichever is less.

If an autopilot is engaged, it will fly the commanded pitch attitude. If in the take-off phase, thrust de-rates are canceled, however, since throttle hold is active, the pilots must manually advance the thrust levers to obtain maximum take-off thrust. If in the approach phase of flight, the autothrottle system advances the thrust levers to "go-around" thrust. The EADI's annunciate "Go-Around" for all

modes, but the commanded pitch attitude is 15o or approximately 1o below the PLI, whichever is less. As the windshear dissipates, the autopilot/flight director system smoothly transitions back to the normal take-off or go-around modes.



WINDSHEAR MODE



EGPWS TERRAIN FUNCTION

Purpose

The GPWS provides terrain awareness displays and alerts. The GPWS can also give alerts and warnings during approach based on terrain clearance.



EGPWS TERRAIN FUNCTION



FMCS - COMMUNICATION, NAVIGATION, SURVEILLANCE

Pegasus FMS Introduction

The Future Air Navigation System (FANS) is a phased improvement of airspace management. ATC has set aside certain oceanic and remote tracks for high density traffic. Operators equipped and certified to improved Communication, Navigation, Surveillance/Air Traffic Management (CNS/ATM) standards may use these tracks. These are the benefits of this system:

• Reduced separation requirements • Flexible tracks • Improved response for altitude and enroute change requests • Avoidance of altitude loss for crossing tracks • Improved availability of alternate airports

Internal Data

Internal data is stored in the FMC. The data is loaded by portable or airborne data loader and may include these:

• The operations program software is the operating system for the FMCS. It defines which sensors are used, how calculations are made, and corrects errors in the steering and thrust commands

• The navigation data base has route structure, airports, way points, nav aids and other important information. This data is updated every 28 days

• The performance data base defines the combination of airframe and engine characteristics in a specific environment. It is part of the FMC as delivered

• The Operational Program Configuration (OPC) data contains Boeing controlled modifiable data. It is separately loaded into the FMCS

• Airline Modifiable Information (AMI) contains data selected by the airline in a separately loaded file

Additional Sensors

The FMC uses GPS satellite data, VOR bearing, DME slant range, and localizer deviation to update position. GPS also gives universal time coordinated for accurate time. Localizer position update is used when these conditions are satisfied:

• ILS procedure is part of the active route • Valid localizer frequency has been tuned for the active runway • Airplane is within 20 NM of runway threshold • Localizer deviation is less than 1.25 dots • Airplane track is within 45 degrees of the runway heading • Altitude is less than 6000 feet

If GPS is available, the FMC uses GPS data to calculate position. If GPS is not available, the primary source of radio position data for the FMC is DME slant range from two nav aids (DME/DME). If DME/DME is not available, the FMC uses VOR bearing and DME slant range from the same nav aid station (VOR/DME). If radio position update data is unavailable, the FMC uses the inertial reference system to calculate position.

MMR’s are installed to replace ILS receivers on GPS equipped aircraft.



FMCS - COMMUNICATION, NAVIGATION, SURVEILLANCE



FLIGHT MANAGEMENT COMPUTER SYSTEM FUNCTIONS

General

The flight management computer system may be used in any of the following configurations:

• FMC/display: A route is selected and activated on the CDU which produces a map display on the EHSI when in the MAP mode. The display may be used for reference while flying the airplane manually.

• FMC/flight director: In addition to activating a route, the flight directors are activated and LNAV/VNAV engaged on the AFCS mode control panel. This enables the crew to fly the selected route following pitch and roll commands displayed on the EADI via the flight director command bars. The FMC provides thrust targets, as seen on EICAS, for the crew's reference.

• FMC/autopilot/autothrottle: In addition to the aforementioned, this mode involves arming the autothrottle and engaging a single autopilot channel to command which enables the FMC to issue steering and thrust commands. This results in aileron movement to track the lateral route profile, and elevator movement coordinated with throttle movement to track the vertical route profile.

CDU/display: The CDU can be used to generate a map display on the EHSI. When the FMC fails, the crew selects the CDU as the source for the map display. The CDU stores the last active route from the FMC and displays the route on the EHSI. The crews can modify the route on the CDU. The display on the EHSI may be used for reference while flying the airplane manually.



FLIGHT MANAGEMENT COMPUTER SYSTEM FUNCTIONS



FLIGHT MANAGEMENT COMPUTER SOURCES

General

The FMC uses GPS satellite data, VOR bearing, DME slant range, and localizer deviation to update position. GPS also gives universal time coordinated for accurate time. Localizer position update is used when these conditions are satisfied:

• ILS procedure is part of the active route • Valid localizer frequency has been tuned for the active runway • Airplane is within 20 NM of runway threshold • Localizer deviation is less than 1.25 dots • Airplane track is within 45 degrees of the runway heading • Altitude is less than 6000 feet

If GPS is available, the FMC uses GPS data to calculate position. If GPS is not available, the primary source of radio position data for the FMC is DME slant range from two nav aids (DME/DME). If DME/DME is not available, the FMC uses VOR bearing and DME slant range from the same nav aid station (VOR/DME). If radio position update data is unavailable, the FMC uses the inertial reference system to calculate position.

MMR’s are installed to replace ILS receivers on GPS equipped aircraft.



FLIGHT MANAGEMENT COMPUTER SOURCES



FLIGHT MANAGEMENT COMPUTER SYSTEM

Introduction

The flight management computer system accomplishes the following:

• Provides a single focal point which enables the crew to select, activate and modify a three dimensional route structure from data stored internally.

• Reduces crew workload by eliminating constant reference to charts and manuals and auto-tunes the necessary radio navigation systems.

• Transmits steering and thrust commands to automatically fly the selected route and displays data for visual monitoring of current dynamic conditions referenced to the route.

Function

The basic function of the FMCS is to compare a selected route (lateral and vertical) to airplane position and use this data to generate steering and thrust requests to maintain the airplane on the requested route profile.

Internal Data

Stored in the FMC, which includes:

Navigation data base - used to define route selection and contains airports, procedures, way points, nav aids etc. This portion of the internal data is inputted and updated by a portable data base loader and connector in the flight deck area. Airline policy file - is part of the navigation data base. The airline policy file contains the required items, airline policy custom file data and the airline policy file options.

External Inputs

CDU - provides the crew interface for inputs to the FMC.

External sensors - provides data to be used for determining the lateral and vertical airplane position.

External Outputs

Guidance commands are translated into control surface movement by the flight control computer (FCC) and throttle lever movement by the thrust management computer (TMC). In addition data to the EFIS symbol generators provides a visual display for crew monitoring.



FLIGHT MANAGEMENT COMPUTER SYSTEM



FLIGHT MANAGEMENT COMPONENT LOCATIONS


Left & right CDU - forward electronics panel (P9)

Data base loader input connector - power distribution panel (P6)

Left FMC, CDU & FMS switching circuit breakers - overhead circuit breaker panel (P11-1)

Right FMC, CDU & FMS switching circuit breakers - overhead circuit breaker panel (P11-4)

FMC annunciator light - captain's instrument panel (P1-3)

Left FMCS source select switch - captain's instrument panel (P1-1)

Right FMCS source select switch - F/O's instrument panel (P3-3)



FLIGHT MANAGEMENT COMPONENT LOCATIONS



FMCS COMPONENT LOCATIONS MEC


Left flight management computer - E2 rack, shelf 2

Right flight management computer - E2 rack, shelf 3

FMC tuning relay, L - Left miscellaneous electric equipment panel (P36)

Left CDU nav enable relay and Left FCC source select relay - E2 rack, shelf 2

FMC tuning relay, R - right miscellaneous electric equipment panel (P37)

Right CDU nav enable relay, right FCC source select relay and Capt. and F/O both on C-SG - E2 rack, shelf 3

Data base loader circuit breaker - forward miscellaneous electric equipment panel (P33)



FLIGHT MANAGEMENT COMPONENT LOCATIONS - MEC



FLIGHT MANAGEMENT COMPUTER

Purpose

The FMC is utilized to translate crew initiated requests and sensor data into maintaining a selected route and furnishing display data for visual monitoring.

Physical Description/Features

Packaging: As depicted in the graphic # Size: 8 MCU (Modular Concept Unit) Weight: 35.5 lb. (16.08 KG) Power: 115 volts ac, 400 Hz, 141.4 watts Cooling: Forced air/ARINC 600

General

The flight management computer (FMC) contains circuitry to support processing of information. The FMC contains 3 16-bit processors that perform mathematical computations and data manipulation; a nonvolatile mass memory system for storage of programs and data bases; an ARINC -429 receiver subsystem for receiving data from interfacing systems and an ARINC -429 transmitter subsystem for transmitting data to interfacing systems.

Front Panel

MAINTENANCE SELF-TEST switch: An INITIATE TEST/LAMP TEST switch initiates the self-test described in the maintenance practices section of this document.

FMC FAIL: A red LED FMC fault annunciator

TEST IN PROCESS: A yellow LED FMC annunciator

Basic Components

Power supply: Single side-mounted, plug-in power supply.

Mass memory storage for the navigation data base.

EEPROM memory storage for:

• Operation program - Performance data base - Nav data base - Guidance buffer (route storage) - Scratch pad memory

Circuit cards: 13 plug-in printed circuit boards

BITE/Monitor

The FMC contains both hardware and software systems which perform a power-up BITE and then immediately shift to a continuous monitor for proper operation.



FLIGHT MANAGEMENT COMPUTER



FMCS - CONTROL DISPLAY UNIT (CDU)

Purpose

The CDU provides the interface between the crew, the FMC and other systems. In the event of FMC failure the CDU can display navigation data (IRS PROGRESS and IRS LEGS pages) and provide map data to the EFIS symbol generators.

Display Format

There are 14 lines of data with a possible 24 characters per line. The top line is the page title and number of pages associated with the display.

Mode Keys

MENU

• Pressing key will display the systems with which the CDU can communicate. Communication between the CDU and various systems can be manually selected

INIT/REF (Initialization/Reference)

• Provides access to pages of data required for initialization of the FMCS and IRS for flight plus APPROACH reference data in-flight

DEP/ARR (Departure/Arrival)

• If no active route has been designated. An index will be displayed. With an active route and on the ground, a departure page from the origin will be shown.

FIX • Allows the creation of waypoint fixes at the intersection of the present route

and selected radials from known waypoints. It is used in conjunction with the EFIS Map display.

HOLD • Provides for definition of a holding pattern at any designated waypoint.

PROG (Progress)

• Displays current dynamic flight and navigation information. Distance to go, ETA and fuel remaining data relating to crew entered alternate destinations can be obtained for comparative purposes.

Function Keys

EXEC (Execute) • The command key of the FMCS. Used for activating the flight plan,

changing the active flight plan or changing the active guidance mode.

CLR (Clear) • Single brief press of key will cause either the last character of a data entry

or a complete message in the scratch pad to be erased. A longer press of key will erase entire data entry.

DEL (Delete) • Pressing of key inserts DELETE into the scratch pad. Line selection

removes data in the associated data field.

CDU Annunciators

MSG (Message) • Illuminates when FMC-generated message is displayed in the scratch

pad.

DSPY (Display) • Illuminates when current display is not related to the active flight plan leg or

the currently operational performance mode.

FAIL • Illuminates when the selected FMC or CDU fails.

OFST (Offset) • Illuminates when a parallel offset is in use.



FMCS - CONTROL DISPLAY UNIT (CDU)



FLIGHT MANAGEMENT EXTERNAL CONTROLS

System Interfaces

The details of the components affecting FMCS computations and/or displays are depicted in the graphic.

General Operation

AFCS Mode Control Panel:

LNAV switch/light - Initiates a request for lateral route navigation to the engaged flight control computer. Acknowledgment is indicated by the switch light bar illuminating. The flight control computers acknowledgment is then sent by the AFCS mode control panel to the flight management computer to complete the request.

VNAV switch/light - Same operation as LNAV for vertical route navigation.

Altitude select control - With the flight being controlled by the FMCS the airplane may not depart from or fly through this selected altitude.

Speed engage/select control - When vertical navigation (VNAV) is operational the speed display on the mode control panel is blank. At this time the FMC determines the airspeed command.

The pilot can manually control the airspeed during VNAV operation by use of the speed select knob. When the speed select knob on the mode control panel is pushed, the speed window unblanks and shows the current airspeed. Rotation of the speed select knob manually enters a new airspeed command. The FMC uses the manually entered command to control the airplane airspeed.

When the pilot pushes the speed select knob a second time, the speed window blanks and the FMC once again determines the airspeed command.

Autothrottle ARM switch - The mode control panel sends an ARM discrete to the thrust management computer which along with valid internal data enables the

FMC to transmit mode and target thrust requests to the thrust management computer.

EFIS Control Panel:

Range switch - The range selection on the EFIS control panel determines how much data is transmitted by the FMC for display on the electronic horizontal situation indicator (EHSI).

Mode switch - Enables the FMC to autotune the VOR/DME’s in MAP and PLAN and utilize FMC data for the EHSI display.

MAP background data switches - Enable the FMC to transmit additional data for display in the MAP mode.

VOR Control Panel:

Enables the crew to interrupt the FMC Autotune capability and provides a display of the tuned navaid frequency.

Thrust Mode Select Panel:

• Allows the crew to select a thrust limit mode. The crew selected mode overrides the FMC selected mode. The TEMP SEL knob allows the crew to select an assumed temperature for takeoff thrust calculations which are performed in the thrust management computer. The CDU displays the assumed temperature on the takeoff reference page.



FLIGHT MANAGEMENT EXTERNAL CONTROLS



FLIGHT MANAGEMENT INPUTS

General Subsystem Features

The flight management computer system uses the navigation, clock, and fuel sensors with an operational program to satisfy crew selected inputs.

General Operation

Fuel Jettison Panel - The FMC senses fuel jettison from the fuel jettison panel. The FMC re-initializes its calculation of total fuel when either nozzle switch is on.

Engine Discretes Cards - The FMC senses the engine bleed status from analog discretes as a secondary source.

Note: The primary source of bleed status is the thrust management computer.

The engine discretes cards provide the following discretes:

• ECS PACK ON • ECS PACK H1 FLOW • ISOLATION VALVE OPEN • COWL ANTI-ICE • WING ANTI-ICE • AIR DRIVEN PUMP ON

Clocks

The FMC uses clock time and date. Once initialized, the FMC uses only minutes and seconds from the clock. Hours are calculated internally. The clock sends day, month and year to the FMC. This input is used to check the active Nav Data Base dates. If the clock date is not within the active Nav Data Base effective dates, an alert message "NAV DATA OUT OF DATE" is displayed in the scratch pad.

EICAS Computer

The FMC uses left and right engine fuel flow from EICAS for its primary fuel quantity computation. If fuel flow is invalid, fuel totalizer is used.

Fuel Quantity Processor

The value of FMC INITIAL FUEL QUANTITY is set equal to the total fuel quantity from the fuel quantity processor when the airplane is on the ground and there is no fuel flow. When fuel flow begins, the initial fuel quantity value is saved for all future calculations and the fuel flow input data from the EICAS computer is integrated, summed and saved. FMC calculated fuel is computed by subtracting the fuel flow sum from the initial fuel quantity value. Throughout the flight, the difference between FMC calculated fuel and total fuel quantity is computed. If the difference exceeds 3000 lb., a CDU alert message is generated (Fuel Disagree - Prog. Pg. 2). To avoid false alarms due to slosh during maneuvers the fuel discrepancy must exceed 2000 lb. for 5 minutes before the 3000 lb. threshold is checked.

Air Data Inertial Reference Unit

Position, velocity, heading, altitude, and vertical speed data are used for navigation and guidance computations.

True airspeed (TAS), computed airspeed (CAS), static air temperature (SAT), and altitude are used for vertical guidance and performance computations.

MMR’s replace ILS receivers and have capability of ILS function as well as GPS. This is for GPS equiped aircraft only.



FLIGHT MANAGEMENT INPUTS



FLIGHT MANAGEMENT OUTPUT DISPLAYS

System Interfaces

The FMCS furnishes data to various components to help the flight crew evaluate flight progress and to assist the maintenance crew in troubleshooting.

EHSI Displays

Two FMCS related display modes are available:

• PLAN - used to examine any segment of the entire route structure.

• MAP - used for inflight monitoring of actual versus selected route profile.

EADI Displays

The FMCS display data is:

• LNAV/VNAV mode annunciation’s • Ground speed digital readout • Selected target speed (when VNAV is operational) • Decision speed, V1 • Rotation speed, VR • Flap maneuver speed EICAS

Thrust target cursor - With VNAV engaged, an FMC computed value of thrust is displayed in magenta to distinguish it from a TMC originated value which would be green.

Messages:

• FMC MESSAGE is an advisory message. FMC MESSAGE occurs when the FMC generates an alert message. You must look at the scratchpad on the CDU to read the alert message.

• L or R FMC FAIL means the left or right FMC has failed. L and R FMC FAIL are advisory messages.

• PILOT RESPONSE means the FMC has not detected activity from the crew for a certain amount of time. The amount of time is set by the airline

in the airline policy file in the FMC. This message may be a warning, a caution or an advisory.

FMC Annunciator Light

Illuminates amber when an FMCS ALERT message is generated in the CDU scratch pad.

Maintenance Control and Display Panel

Stores FMCS fault data accumulated during a flight. This information is utilized to analyze, isolate and correct problems.

RDMI - When the scanning DME option has been selected via the airline policy file and the EFIS control panel is in MAP or PLAN mode, the FMC will direct the RDMI’s to show DME distance from the FMC.



FLIGHT MANAGEMENT OUTPUT DISPLAYS



FMCS - DUAL SYSTEM COMMUNICATION

General

The flight management computer system has two flight management computers (FMC) and two control display units (CDU). Any FMC/CDU pair make an operational system.

Both CDU’s operate alone and at the same time, but are normally updated with display data from the on side FMC.

When a change is made on one CDU it is shown on the other CDU if the same page is displayed.

FMC/CDU

A request from either CDU is completed by the master FMC first and then by the slave FMC.

FMC/FMC

The FMC’s communicate via the intersystem bus. With the initial application of power, the operational program, performance data base, program pins and the navigation data base are compared. If a difference occurs, the fault displays are shown.

FMC/FCC and TMC

The master FMC sends thrust limit mode requests, autothrottle mode requests, and speed or thrust targets to the thrust management computer. The master FMC also provides lateral and vertical steering commands to the flight control computers.

The master FMC is determined as follows:

• Flight director only status - the captain's selected FMC is the master when the captain's F/D switch is on. The first officer's selected FMC is the master when the first officer's F/D switch is on and the captain's F/D switch is off.

• Command status - the captain's selected FMC is the master when the left or center FCC is engaged to CMD. The first officers selected FMC is the master when the right FCC is engaged to CMD.

FMC/EFIS

Each FMC provides the captain's or first officer's electronic flight instrument map displays on the electronic horizontal situation indicator (EHSI). Normally the left FMC provides the display for the captain and the right FMC provides the display for the first officer via the left and right EFIS symbol generators.



FMCS - DUAL SYSTEM COMMUNICATION



FLIGHT MANAGMENT (CDU OUTPUTS)

System Interfaces

Each CDU has two output busses as shown in the graphic.

General Operation

A. Subsystem output bus

Each CDU sends messages to other subsystems via a single arinc 429 data bus. The CDU transmits a unique address label with the data. This label identifies which subsystem the CDU has selected for communication. Thus, each subsystem selects only the data which carries its unique address label.

B. Standby Nav (EFIS output bus)

Each CDU can send data to EFIS in order to provide a map display. The left CDU can send data to the left or center symbol generator. The right CDU can send data to the right or center symbol generator. When the FMC fails, the crew selects the CDU as the source for map display. The CDU stores the last active route from the FMC and displays the route on the EHSI. The crew can modify the route on the CDU. The display on the EHSI may be used for reference while flying the airplane manually.



FLIGHT MANAGMENT (CDU OUTPUTS)



FMCS EFIS DISPLAY

EHSI Map Mode

The FMC provides map background data, distance to go to the active waypoint, ETA at the active waypoint, airplane present position, wind speed and direction, and vertical deviation (not shown).

EHSI Plan Mode

The FMC provides a map display reference to true north and also distance to go, and ETA at active waypoint.



FMCS EFIS DISPLAY



FLIGHT MANAGEMENT CDU MESSAGES

Operation

Alerting and advisory messages are generated by FMC software when a condition exists which degrades the operational viability of the system.

All alerting and advisory messages illuminate the message (MSG) annunciator light on the CDU.

Only the alerting messages set the CDU MESSAGE output discrete which causes:

• The upper EICAS display to display FMC MESSAGE (Level C)

• The FMC annunciator light (P1) to illuminate

Control Sequence

Messages/data are assigned a priority as follows:

Priority 1 ALERT MESSAGES 2 ENTRY ERROR ADVISORY MESSAGES 3 ALPHA NUMERIC DATA 4 ADVISORY MESSAGES In addition to a priority by category, a chronological priority is assigned to each category with the highest priority assigned to the most recent data.

Un-cleared messages and un-cleared alpha-numeric data are stored in a message stack. As alerting messages are generated, they are displayed in the scratch pad of each operating CDU. As uncleared alerting messages are pushed down by other alerting messages, they are sequentially added to the top of the stack.

As the CLR key is pressed, in discrete steps, the stack is displayed and messages and data cleared sequentially from the top to the bottom (holding the CLR key down shall not cause all messages and data to be displayed and cleared in a continuous sequence).

ALERT Messages

Alerting messages appear in the scratch pad of each operating CDU regardless of the prior contents of the line.

ENTRY ERROR ADVISORY Messages

Advisory messages displayed as a result of data entry errors have a higher priority than data or other advisory messages displayed in the scratch pad.

As data entry error advisory messages are generated, they are displayed in the scratch pad of the CDU on which the data entry attempt was made. However, if an alerting message or messages are required to be displayed, the data entry error advisory message or messages are inserted below the alerting messages in the stack.

ADVISORY Messages

As advisory messages are generated they are displayed in the scratch pad if an alerting message or a data entry error advisory message or data is not displayed, otherwise they are inserted below alerting messages and data entry error advisory messages and data in the stack. These advisory messages are displayed in the scratch pad of each operating CDU except where the set logic requires a particular CDU page be accessed in which case they are displayed only on the CDU with the required page displayed.

The MCDU was designed as a stand alone item, therefore it can generate messages that relate to the MCDU's independent operation.



FLIGHT MANAGEMENT CDU MESSAGES



MAINTENANCE INDEX PAGE

Purpose

This graphic shows the maintenance pages that are available, on the ground, to assist in isolating problems with the flight management computer system. All maintenance pages are accessed from a single maintenance index.

Page Access

The maintenance index may be accessed from the INIT REF INDEX (initialization/reference index) when the airplane is on the ground.

Data Fields

1L CROSS LOAD - used for transferring the navigation data base between the FMC’s

2L IRS MONITOR - shows inertial reference system (IRS) position error rate at flight completion for each inertial reference unit

3L SENSORS - shows the current status for sensors which provide data to each FMC

4L DISCRETES - shows status of analog discretes to each FMC

6L INDEX - returns display to the INIT/REF INDEX

6R POLICY - shows the performance factors that are part of the airline policy file in the nav data base.



MAINTENANCE INDEX PAGE



NAV DATA CROSSLOAD PAGE

Page Access and Purpose

The NAV DATA CROSSLOAD page is accessed as indicated on the graphic and enables one FMC to transfer its navigation data base to the other FMC.

NAV DATA CROSSLOAD Page Data Fields

2L Navigation data base identifier as displayed on line 2L of the IDENT page.

3L-3R TRANSMIT, RECEIVE - Data base crossload is initiated by selection of field 3L on one CDU followed by 3R on the other. For proper transfer to take place, this must be preceded by enabling the NAV DATA UPDATE input discrete on both systems by inserting "ARM" into the scratchpad and pressing LSK 6R. Selection of TRANSMIT (3L) deletes RECEIVE (3R), and vice versa, on the same CDU.

4L-4R Prior to selection, line 4 is blank. Following cross load selection TRANSFER IN PROGRESS appears. Upon completion, TRANSFER COMPLETE appears. If the transfer is unsuccessful, TRANSFER ABORTED appears in line 4.

6L Accesses the MAINT INDEX page.

6R Selection of ARM is required to enable page function. Leaving the page deletes ARM.



NAV DATA CROSSLOAD PAGE



IRS MONITOR PAGE


The IRS MONITOR page is accessed as indicated on the graphic and enables an evaluation of position error for each inertial reference unit.

IRS Monitor Page Data Fields

2L Position error rate for each 3L IRS, computed by dividing the 4L distance from the FMC computed to the IRS position by the total flight time.

These values are computed at flight completion and cleared when airborne or at power down.

6L Returns the MAINT INDEX display.



IRS MONITOR PAGE



REFERENCE SENSOR STATUS PAGE


The L/R FMC SENSOR STATUS pages are accessed as indicated on the graphic and shows real time status of all sensors associated with either the left or right FMCS.

SENSOR STATUS Page Data Fields

One of four displays is associated with each sensor:

• OK Connected and functioning properly • FAIL Not connected (or) No power (or) Not functioning correctly • TEST In SELF-TEST mode • Unit not required for this FMCS system or for an airplane configuration.

6L Accesses the MAINT INDEX page.



REFERENCE SENSOR STATUS PAGE



IDENTIFICATION PAGE

Purpose

Provides a means of reviewing the FMC nav data base and program configuration.

Page Access

FMC line selection on the MENU page.

IDENT line selection on the INIT/REF INDEX page.

1L-MODEL NUMBERDisplays the aircraft model as read from the engine/airframe identification pins. If the identification pins do not match the stored performance data base, then blanks will be displayed.

2L-NAV DATADisplays the data base identifier in large font. If the data load complete bit is not set, this field will be blank.

4L-OPERATIONAL PROGRAM NUMBERDisplays the operational program part number.

5L-DRAG/FFFuel mileage factor assigned to drag and fuel flow computations are expressed as percentages. No 5L entries allowed in the air. On the ground, drag and fuel flow factors will not be enter able until the field has been armed by entering ARM into 5L. At this time, ARM will be displayed in small font just to the right of DRAG/FF. The field will remain armed until leaving the IDENT page. Once armed, entry rules are as follows:

• Valid entry range is from -5.0 to +9.9. • Entries are maintained over long term power interrupts and flight

completion. • If no value has been entered, +0.0/+0.0 will be displayed. • Fuel flow only entry requires a leading slash, whereas the drag factor has

optional slash entry.

6L-INDEXSelects the INDEX page.

1R-ENGINE IDENTIFICATIONDisplays the engine identification number as read from the engine/airframe identification pins. If the identification pins do not match the stored performance data base, then blanks will be displayed.

2R/3R-NAV DATA BASE EFFECTIVITYThese two lines can be interchangeable via line selection for the purpose of activating a new nav data base or recalling the old one. Selection can only be on the ground and an entry into 2R clears out any previously selected flight plan. The active nav. data base calendar cycle is monitored by the FMC and is checked against the source clock's calendar date. If the clock's calendar date exceeds the active nav data base calendar cycle, the FMC will generate a NAV DATA OUT OF DATE message.

4R-VERSIONDisplays the version identification for the current configuration.

5R-CO DATADisplays the airline policy file identifier in large font. If the nav data base load complete bit is not set, this field will be blank.

6R-POS INTSelection displays POS. INT page.



IDENTIFICATION PAGE



AIRLINE POLICY PAGE

Purpose

From the FMS CDU there are several pages that assist in maintenance of the FMS system. One of these is the Airline Policy Page. This page can be reached by selecting “POLICY” from the Maintenance Index page. The Airline Policy Page gives specific Company option codes, margins, accelerations and thrust reductions.



AIRLINE POLICY PAGE



FMCS SELF TEST

General

The FMCS maintenance self-test is activated on the ground by pushing the "INITIATE TEST/LAMP TEST" button on the front panel of the FMC. While the button is pressed, the red FAIL LED comes on. Once the test is initiated, the yellow TEST IN PROCESS LED comes on and stays on until the test is finished. The processor interrupts all subsystems and places them in a comprehensive self-test mode. Upon successful completion of the test, the red and yellow LED’s go off.

Self-Test Annunciations

Displays associated with a successful and a "failed" self-test result are shown on the graphic.



FMCS SELF TEST

B767 ATA 34 Student Book

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Transcript of B767 ATA 34 Student Book