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B767sim

B767simulator

Avionics bending: Symbol Generator

A. The flight instrument system provides the main displays for most of the airplane's navigation systems. It includes the electronic flight instrument system (EFIS), the radio distance magnetic indicators (RDMI), and the vertical speed indicators (VSI). It also includes the switches on the instrument source select panels.

B. The EFIS uses CRT indicators. It provides multicolor navigation displays. The system displays the following:

(1) Pitch, roll, and directional data; map displays and flight path data; weather radar data; altitude and decision height; autopilot mode data; and input system fault annunciations.

(2)Airspeed data.

(3)Traffic alert and collision avoidance data.

C. The two displays associated with the EFIS are the Electronic Horizontal Situation Indicator (EHSI) and the Electronic Attitude Director Indicator (EADI). Two units of each display are installed.

Each set operates independently under normal conditions. Each display set has a dedicated control panel, remote light sensor, and a symbol generator. A third symbol generator is installed, as a backup, for either display set.

D. The switches on the instrument source select panels provide selection for backup data sources in the event of a main data source failure.

E. Normally, the left and right symbol generators supply the corresponding display set. If a fault is detected by the operating symbol generator, either display set can be switched to the center symbol generator. The RDMIs and VSIs can be switched to the center IRU if a source fault is detected.

 

F. If the captain and F/O select the center EFIS system, a level B caution message - INSTR SWITCH - will appear on the EICAS Display.

 

G. The RDMIs are the secondary heading, bearing, and distance displays. A portion of the indicators display the airplane's present heading. They also display the directional bearing and navigational distance to selected reference points.

H. The RDMIs receive this data from the IRS, VOR, DME, and ADF systems.

I. The VSIs are the main vertical rate-of-climb and descent display. They receive this data from the inertial reference system.

 

 

 

In the following pics, a view inside of a Symbol generator during cleaning process

 

 

All Boars and power supply removed

 

Power supply installed

 Display driver card installed

 All cards installed, ready to be closed

SG front view

 

 

Avionics bending: P51 Warning System

1. General

A. The warning system provides the flight crew with visual and aural warnings. Warnings are provided for abnormal airplane system conditions. They are also provided for airplane out of configuration conditions for various operating modes.

B. The warning system consists of a warning electronics unit, aural warning loudspeakers, aural cancel switch, and test switches. It also includes the master WARNING switch-lights, SPEEDBRAKES light, and discrete warning light.

C. Input signals are received from airplane sensors, other airplane systems, or generated by the pilots. Signals are processed within the warning electronics unit (WEU).

D. The visual displays are level A warning messages on the display unit and red WARNING lights. The aural signals are bell, siren, owl, and ground proximity warning system voice messages.

 

2. Component Details

A. Warning Electronics Unit (WEU)

(1) The warning electronics unit (WEU) P51 is located in the forward electrical/electronic equipment area.

(2) The WEU contains two power supply modules with fault indicators and reset switch, and plug-in modules which generate warning signals. The WEU provides output signals to the warning loudspeakers, master warning lights, and discrete warning lights. The unit weighs 33 lbs and is cooled by externally blown air.

(3) The modules located in the WEU and their functions are as follows:

 (a) Power Supply Modules

1) The two redundant power supplies receive 115 volts ac power and develop the +5 volts and ±12 volts dc power required for all warning modules. There are two ball type fault indicators, one for each module. Each ball latches to black when the corresponding power supply is operating properly and latches to yellow when a fault occurs. The RESET switch is provided to set the fault balls back to the operational mode.

2) The two power supply fault indicators and a reset switch are located on the WEU BITE module inside P51.

 (b) EMI Filter Module

 

1) There are two EMI filter modules in the WEU. Each filter module is paired with a power supply module. The EMI filter module connects to the power line and filters the AC line noise.

 

 (c) Stall Warning Module

 

1) There are two stall warning modules. The left module provides the captain's warning signal and the right module, the first officer's warning signal. The modules are each microprocessor controlled.

 

 (d) WEU BITE Module

 

1) The WEU BITE module provides fault indications for the two power supplies and the two stall warning computers.
 

 (e) Altitude Alert Module
 

1) The altitude alert module is microprocessor controlled. It provides altitude acquisition advisory and deviation cautions.

 (f) Takeoff Configuration Warning Module

1) The takeoff configuration warning module provides warning signals for improper airplane configuration during takeoff.

 (g) Landing Configuration Warning Module

1) The landing configuration warning module is microprocessor controlled. It provides warning signals for improper airplane configuration during landing and speed brake caution signals.

 (h) Master Warning Module

1) The master warning module is microprocessor controlled and provides the warning signals to turn on the master warning lights.

 (i) Bell/Chime Aural Warning Module

1) The bell/chime aural warning module provides fire bell warning and crew call chime signals to the audio amplifiers in the siren/owl aural warning modules.

 (j) Siren/Owl Aural Warning Module

1) The left and right siren/owl aural warning modules supply level A warning and level B caution signals.

2) The siren warning is activated by overspeed, cabin altitude exceedance, T/O configuration warnings, and Landing configuration warnings.

3) The siren warning is also activated by autopilot disconnect.

4) The owl warning is activated by level B EICAS messages and altitude alert.

5) The siren/owl modules also supply amplification of the bell/chime warnings and ground proximity voice.

6) The siren/owl modules also amplify voice messages received from the TCAS.

 (k) Clacker/Wailer Aural Warning Module

1) The clacker/wailer aural warning modules provide autopilot warning signals to the siren/owl aural modules. The clacker is not used.

 (l) EICAS Signal Consolidation Card (SCC)
 

1) The EICAS SCC converts analog signals to digital signals which are then sent to the EICAS function.

 

B. Configuration Warning Test Switch

(1) The configuration warning test switch is located on the miscellaneous test panel on the right side panel. It is a two-position, spring loaded switch. The T/O position is for testing the takeoff configuration warning circuits. The LDG position is for testing the landing configuration warning circuits.

C. Gnd Prox/Config Gear Ovrd Switch

(1) The Gnd Prox/Conf Gear Ovrd Switch is located on the first officer's instrument panel. It is an alternate action light switch. The switch cancels the landing configuration aural warning and displays the amber OVRD message when pressed.

D. Master Warning Lights

(1) The red master WARNING lights are located on both ends of the pilots' glareshield P7. The lights come on when the master warning module in the warning electronics unit generates a warning signal. Pressing the indicator light cap will turn off the light and silence the fire bell warnings, Landing configuration warning, or the cabin altitude warning.

E. Discrete Warning Light

(1) The red CONFIG discrete warning light is located on the center instrument panel. The light comes on when the configuration warning module in the warning electronics unit generates a warning signal. The light is powered by the master dim and test circuit.

F. Loudspeakers
 

(1) The warning loudspeakers are located above the captain's and first officer's seats. Each loudspeaker is a permanent magnet speaker driven by an internal single stage amplifier. The two speakers provide aural tones and messages when turned on by either the left or right siren/owl aural warning module.

G. Speed Brake Handle Position Switch

(1) The speed brake handle position switch is located on the control stand. The switch is a thru-contact microswitch with a roller control arm. If the switch is set at the down detent position, a ground is applied to the takeoff configuration warning module. A warning is given when the switch is not in the down detent.  

 

 

 

 

 

 

In the following video a power up test and an aural test using a loudspeaker from cockpit.

 

 

 

 

 

 

 

 

Avionics bending: CVR

1. General

A. The voice recorder system preserves a continuing record of the latest 30 or 120 minutes of flight crew communications and conversations. The four channel voice recorder receives inputs from the audio selector panels for the captain, the first officer, and the first observer, and from an area microphone in the flight compartment.

 

B. The voice recorder system consists of a recorder located in the aft passenger cabin ceiling and a control panel with an area microphone located on overhead panel P5.

C. 115 VAC power is supplied to the system from the right AC bus. The VOICE RECORDER circuit breaker, located on overhead panel P11, controls power to the system.

2. Component Details

A. Voice Recorder

(1)

The voice recorder is installed in the E7 equipment rack in the aft passenger cabin ceiling. The recorder is a thermally insulated, impact resistant, 4 MCU unit, and international orangein color.

(2) 

The voice recorder contains a tape transport which includes the tape, tape drive, two four-track recording heads, two erase heads, two monitor heads, and a bulk erase coil. The unit also contains four recording amplifiers, a monitor amplifier, 65 kHz bias generator,
600 Hz test circuits and power supply.

(a) The recording tape is 308 feet long, 1/4 inch wide in a continuous endless loop. The four track tape runs at 1 7/8 IPS providing 30 minutes of recording time.
(b) The front panel contains a headphone jack and a underwater locator beacon (ULB). A headset inserted into the headphone jack will monitor all channels slightly delayed after they are recorded.
(3) An ULB is attached to the voice recorder front panel and is a self-contained device. The ULB will emit a 37.5-KHz signal when the water activated switch is closed. The ULB contains a dated label to indicate when the battery requires periodic replacement.

 

B. Voice Recorder Control Panel

(1)

The voice recorder control panel is located on overhead panel P5.

It contains an area microphone with preamplifier and filter circuits, a HEADPHONE jack to monitor system recording, a TEST switch to initiate the functional self test, a light or meter to
provide an indication of self test success or failure, and anERASE switch for data erasure.

(2)

Push the TEST switch will activate test circuits within the voice recorder. A tone is applied to all four channels. If the test is successful, the needle on the monitor meter moves and stays in the green area and you hear a tone in the headse

 

(3)

Push the ERASE switch will cause the voice recorder data to be completely erased. Bulk erasure is possible only when the airplane is on the ground with the parking brake set.

 

   
  

 

 

Avionics bending: MCDP computer

A.

The maintenance monitor system consists of a Maintenance Control Display Panel (MCDP) that monitors the status of the Flight Control Computers (FCC), Flight Management Computers (FMC), Thrust Management Computer (TMC), and their related sensors. The MCDP interrogates the computers after each landing, and stores up to five flight faults per computer in non-volatile memory for later interrogation by maintenance personnel.

B.

A central processing unit in the MCDP controls and processes ground test commands, ground test functions, and on-ground and in-flight faults. The MCDP displays flight and ground faults, test number, name or status, and operator instructions. The MCDP controls and display panel are located on the front of the unit. The controls enable maintenance personnel to display faults and ground tests that include operator instructions. The MCDP is located in the main Electrical/Electronic (E/E) equipment center.

C.

The remote control panel is mounted in the P61 panel and is used for displaying flight faults and for running ground tests from the flight deck. The MCDP remote display is through the EICAS system maintenance panel in the P61 panel. The MCDP display can be displayed on the EICAS with or without the remote control panel, by pressing the CONF/MCDP switch of the EICAS maintenance panel.

D.

Maintenance Control Display Panel Interfacing Systems

 

(1) The primary MCDP interfaces are with each FCC, FMC, and TMC.Secondary interfaces with the MCDP consist of FCC, FMC, and TMC individual system sensors. Both primary and secondary interfaces are monitored by the MCDP during ground test functions.

(2) Primary Interfaces

(a) The left, center, and right flight control computers use ARINC 429 data buses for transmitting interface fault data to the MCDP and receiving ground test data from the MCDP. Analog discrete lines are used for ground test control.

(b) The left and right flight management computers use ARINC 429 data buses to transmit interface fault data to the MCDP. The MCDP can only interrogate the FMC for fault data. The MCDP ground tests do not affect the FMC systems.

(c) The thrust management computer uses an ARINC 429 data bus for transmitting interface fault data to the MCDP and receiving ground test data from the MCDP. An analog discrete line is used for ground test control of the TMC by the MCDP.

(3) Secondary Interfaces

(a) The mode control panel supplies control signals and receives status data from each FCC, FMC, and TMC on ARINC 429 data buses.

(b) Each FCC supplies analog control signals to its control servos and receives analog servo position signals. Each FCC receives sensor data on ARINC 429 data buses from the

associated Inertial Reference Unit (IRU) and air data computer.

(c) The TMC receives control signals on an ARINC 429 data bus from the thrust mode select panel. The TMC also transmits status data to the thrust mode select panel on an ARINC 429 data bus. The TMC supplies analog control signals to and receives analog

position signals from the autothrottle servomotor generator.

 

 

 

 

 

 

 

 

Avionics bending: Tape Radio Altimeter



 

A. The radio altimeter (RA) system supplies vertical position data for use by the pilots for runway approach, landing, and takeoff. The RA system provides accurate measurement of absolute altitude (height above terrain) from 2500 ft. to touchdown. Altitude data is routed to user systems on 429 digital data buses.

B. Three complete systems are installed. Each system consists of a receiver/transmitter (R/T), and one transmit and one receive antenna. Altitude data is displayed on the electronic attitude director indicators (EADIs)and on tape radio altitude indicator if equipped.

Avionics bending: IRS Mode Select Panel

The IRS mode select panel (IRMP) is used to control the three IRUs. It provides individual mode selection, displays align status, displays fault annunciation for each IRU, and provides a display and keyboard for IRU initialization and data display. The IRMP is located on the pilot's overhead panel.

The IRMP has three rotary switches for selecting the following modes of each IRU.

(a) OFF - IRS is off.

 

(b) ALIGN (alignment) - During the ALIGN mode, the IRU is initialized. The position information given to the IRS is checked for accuracy in the IRU. This process, when properly completed, allows the IRU to advance to the NAV mode.

(c) NAV (navigation) - The NAV mode is the normal operating mode for the IRS. In this mode, the IRS performs inertial navigation functions and outputs normal IRS data to be

displayed or used by other systems.

(d) ATT (attitude) - The ATT mode is used when failure or total power loss (AC and DC power) is detected in the NAV mode. In this mode, only attitude data is output to the user systems.

The ALIGN, NAV, ATT, and OFF modes are entered using the following mode sequences. The corresponding IRU operation and ALIGN annunciator status are as follows:

################################################################################

| MODE SEQUENCE | IRU OPERATION | ALIGN LIGHT |

################################################################################

|OFF TO ALIGN | REMAINS IN ALIGN MODE AND CONTINUES FINE| REMAINS ON |

| *[1] | LEVELING AND EARTH RATE ESTIMATION. | |

| | | |

|OFF TO ALIGN TO NAV| PERFORMS 10-MINUTE ALIGNMENT AND | ON DURING |

| *[1] *[2] | AUTOMATICALLY SEQUENCES TO NAV MODE. | ALIGNMENT |

| | | |

|OFF OR NAV TO ATT | FULL SUPPORT OF ATTITUDE - RELATED AND | ON FOR 30 |

| | BODY ACCELERATION OUTPUTS IN 30 SECONDS.| SECONDS |

| | | |

|NAV TO ALIGN | SETS VELOCITIES TO ZERO AND CONTINUES | REMAINS ON, |

| *[1] *[3] | TO REFINE LEVELING AND HEADING | FLASHES IF |

| | DETERMINATION (NAV MAY BE RE-ENTERED | MOVEMENT |

| | IN APPROXIMATELY 30 SECONDS). | |

| | | |

|NAV TO ALIGN TO NAV| SETS VELOCITIES TO ZERO. | ON FOR 30 SEC, |

| *[1] *[3] | | FLASHES IF MOV-|

| | | EMENT WHILE ON |

| | | |

|ATT TO NAV OR ALIGN| REMAINS IN ATT MODE UNTIL SWITCHED OFF. | OFF |

| | | |

|ALIGN TO NAV | IF ALIGN TIME GREATER THAN 10 MINUTES, | ON DURING |

| | IRU SEQUENCES TO NAV, HOWEVER, AN ALIGN | ALIGNMENT |

| | TIME OF 17 MINUTES IS REQUIRED ABOVE | |

| | 70 DEGREES NORTH LATITUDE. | |

| | | |

|ATT OR NAV OR ALIGN| 30 SECOND POWER-OFF COUNTDOWN, BITE | ON FOR 30 |

|TO OFF | INFORMATION, PPOS, AND AUTOCAL TRANSFER | SECONDS |

| | TO NVM. | |

################################################################################

 

*[1] Details of mode sequences to ALIGN mode are described in Functional Description of IRS Alignment.

*[2] The mode sequence "OFF TO ALIGN TO NAV" should not be used above 70 degrees North latitude. A full 17 minute alignment is required.

*[3] The IRU should not be sequenced out of NAV mode above 70 degrees North latitude. Accuracy updates require a full 17 minute alignment.

Each mode select switch has a detented NAV position which prevents accidental movement of the switch out of the NAV mode. When the switch is in the NAV position, it must first be pulled out of detent before selecting a new position to prevent damage to the switch.

Three sets of four lights provide system status and fault indication. ALIGN (white when lit), ON DC, DC FAIL, and FAULT (amber when lit) lights are provided for each IRU. These lights work as follows:

(a) The ALIGN light denotes that the IRU is in the align mode and is running an initial position determination, is in initial attitude mode, or is powering down.

(b) The ON DC annunciator lights when the IRU has switched to backup battery power.

(c) If battery power fails, the DC FAIL light will come on.

(d) A lit FAULT annunciator indicates that a BITE detected failure has occurred. The FAULT light also comes on if certain alignment tests fail. These tests are covered in detail in the

operation section.

IRMP Keyboard and Display

(a) The SYS DSPL switch selects the IRU, for on-line interface with the IRMP. The IRMP can only display data from the IRU which has been selected by the SYS DSPL switch.

(b) The DSPL SEL switch selects the type of data to be displayed on the IRMP numeric displays. The IRU, as selected by the SYS DSPL switch, supplies the data. The four switch positions and the data displayed for each position is as follows:

1) TK/GS - Track angle (TK) is displayed in the left display and the ground speed (GS) in the right display.

2) PPOS - Latitude is displayed in the left display and the longitude in the right display.

3) WIND - Wind angle is displayed in the left display and the wind velocity in the right display.

4) HDG - True heading is displayed in the left display and the right display remains blank.

(c) There are two numeric displays on the IRMP. When the IRMP is receiving IRU data, the DSPL SEL switch determines the data on the display. When the IRMP keyboard is used to initialize an IRU, the data punched in at the keyboard is shown on the two displays. For invalid data from an IRU, both displays are blanked. A brightness control for the displays is located concentric within the DSPL SEL switch.

(d) The keyboard consists of twelve lighted keys. To change the numeric display from the IRU receive mode to a keyboard display mode, one of the following keys must first be pressed: N(2), S(8), H(5), E(6), or W(4). Any other initial key is ignored.

1) Pressing N(2) or S(8) once will cause a N or S to appear in the left display. These represent north and south and are used to initialize latitude in the IRU.

2) Pressing W(4) or E(6) once will cause a W or E to appear in the right display. These represent west and east and are used to initialize longitude in the IRU.

3) Pressing H(5) once will also switch the IRMP from an IRU receive mode to the keyboard display mode. This is used to enter magnetic heading in the ATT mode.

4) Numeric data can be entered after one of the five letter keys is pressed. It will be appropriately displayed in the numeric display as it is entered.

5) Keys with letters on them, as well as numbers, provide the letter value when they are the first punched key in a program sequence.

6) The ENT key, when pressed, transfers the data in the IRMP to the IRU. Also, when the ENT key is pressed, the display is first blanked and then returned to the IRU receive mode.

7) The CLR key clears the display, then returns the IRMP to the IRU receive mode.

The IRMP also has a Time to NAV (TTN) display and a maintenance fault code display.

(a) The Time to NAV (TTN) feature counts down the minutes to alignment completion.

(b) The IRU must be in the ALIGN mode, the IRMP mode select switch must be set to ALIGN or NAV, and the DSPL SEL switch must be set to HDG. A single digit will appear on the right side of the IRMP display to indicate minutes remaining as follows:

#########################

| MINUTES | NUMBER |

| REMAINING | DISPLAYED |

#########################

| 10 | 7 |

| 9 | 7 |

| 8 | 7 |

| 7 | 7 |

| 6 | 6 |

| 5 | 5 |

| 4 | 4 |

| 3 | 3 |

| 2 | 2 |

| 1 | 1 |

| 0 | 0 |

#########################

1) The TTN zero will not be visible with the mode select switch set to NAV. The display will blank as the IRS sequences to NAV and the ALIGN annunciator goes off.

2) When realign mode is selected, the TTN will display 7 and count toward 0 in 30 seconds.

 

(d) The maintenance fault code display mode provides access to some IRU faults without external test equipment. Faults are classified as either critical, noncritical, or BITE memory

only. The maintenance fault code display mode will display critical and noncritical faults, but external test equipment is required to read faults stored in BITE memory. Regardless of fault classification, all faults are stored in BITE memory when the IRU is turned off.

1) To enter the maintenance code display mode, first select the desired IRU using the SYS DSPL Select Switch. Next, set the DSPL SEL switch to the HDG position. Enter a 0 on the

keyboard, followed within five seconds by a 1. The highest priority code will be displayed in the two right digits of the IRMP display. Record the code, press CLR to display the next priority code, and repeat to cycle through the remaining codes.

2) If the IRS was navigating, the last displayed value of true heading will be frozen on the display.

3) If TTN display was present, the display will be frozen when the maintenance code is entered. The internal clock will continue to run and a new TTN value will appear after all

maintenance codes are displayed.

4) The following table indicates the corresponding fault for each IRMP code:

###########################################

| IRMP CODE AND | |

| PRIORITY | IRU FAULT |

###########################################

| 01 | POWER SUPPLY CRITICAL |

| 02 | DIGITAL I/O WRAP-AROUND |

| 03 | RAM/NVM/PROM MEMORY |

| 04 | LSIC |

| 05 | DISCRETE INPUT |

| 06 | PROCESSOR |

| 07 | GYRO |

| 08 | ALIGN/SYSTEM |

| 09 | A/D MUX DATA TRANSFER |

| 11 | POWER SUPPLY |

| 12 | ADC DISCRETE OUT |

| 13 | NOT USED |

| 14 | CALIBRATION PROM |

| 15 | INA/OTA |

| 16 | ANALOG PITCH RATE |

| 17 | GYRO |

| 18 | GYRO CONFIGURATION |

| 19 | TEMPERATURE SENSOR |

| 20 | NOT USED |

###########################################

6) Codes 01 thru 09 are critical faults. A critical fault turns on the IRMP FAULT light.

7) Codes 11 thru 20 are noncritical faults and are indicated as follows:

a) During initialization, align, align downmode, or navigate mode, if a non-critical fault is detected on the ground, the IRMP fault annunciator turns on. If a non-critical fault is detected in the air, the IRMP fault annunciator turns on below a set ground speed after touchdown.

b) During attitude mode,no failure indication is given by the IRMP fault annunciator. A fault detected prior to entering

attitude mode is indicated as previously described. Entering attitude mode causes the IRMP fault annunciator to go out.

 

 

 

 

Here are few pictures, a video of light test and manual keyboard position immission and a video with demonstration of ARINC data feeded to the unit's left channel.