ATA 31 - SEP 2024 De-Havilland DHC-8-400 PDF

COURSE CODE DGM12CBXX3Q4PW De-Havilland DHC-8-400 B1 & B2 TRAINING MANUAL ISSUE 02 DATE 20 Sep 2024 REVISION DATE...

COURSE CODE DGM12CBXX3Q4PW De-Havilland DHC-8-400 B1 & B2 TRAINING MANUAL ISSUE 02 DATE 20 Sep 2024 REVISION DATE CHAPTER ATA 31 INDICATING AND RECORDING SYSTEM THIS MANUAL IS INTENDED FOR TRAINING PURPOSE ONLY REV 2.0 20 Sep 2024 PAGE 1 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL THIS PAGE IS INTENTIONALLY LEFT BLANK REV 2.0 20 Sep 2024 PAGE 2 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION CONTENTS DHC-8-400 B1 & B2 TRAINING MANUAL CONTENTS CONTENTS............................................................................................................................................................................................................................... 3 INDICATING AND RECORDING SYSTEMS, GENERAL................................................................................................................................................... 4 ENGINE COCKPIT INTERFACE UNIT................................................................................................................................................................................. 7 INDEPENDENT INSTRUMENTS......................................................................................................................................................................................... 12 RECORDERS.......................................................................................................................................................................................................................... 21 SIGNAL CONDITIONING UNIT.......................................................................................................................................................................................... 37 EXTENDED−STORAGE QUICK ACCESS RECORDER SYSTEM (EQAR).................................................................... Error! Bookmark not defined. MICRO QUICK ACCESS RECORDER SYSTEM (MQAR)................................................................................................ Error! Bookmark not defined. CENTRAL COMPUTER......................................................................................................................................................................................................... 39 CENTRAL WARNING SYSTEM........................................................................................................................................................................................ 101 CAUTION AND WARNING LIGHTS SYSTEM................................................................................................................................................................ 103 TAKE−OFF WARNING SYSTEM....................................................................................................................................................................................... 125 ELECTRONIC INSTRUMENTS SYSTEM......................................................................................................................................................................... 128 REV 2.0 20 Sep 2024 PAGE 3 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION INDICATING AND RECORDING SYSTEMS, GENERAL DHC-8-400 B1 & B2 TRAINING MANUAL INDICATING AND RECORDING SYSTEMS, GENERAL Clocks Introduction The electronic clock has a quartz time base that supplies a continuous display of Greenwich Mean Time (GMT) or Local Time (LOC). The electronic clock can The indicating and recording system has different sub−systems to do the also be set to show the Elapsed Time (ET), the date or set to the chronometer functions that follows: function (CHR).  Show time Flight Data Recorder System (FDR)  Records aircraft parameter data The Solid State Flight Data Recorder (SSFDR) records aircraft parameter data  Receive data from sensors and avionics systems and supply it to other and stores it in crash−protected memory for future retrieval purposes. systems  Make warning tones Central Computer  Show caution and warning lights  Show advisory lights The central computer has a Flight Data Processing System (FDPS) to receive  Take−off warning calculation data from sensors and avionics systems and supply it to other systems. The  Show navigation, engine, and system parameters. Flight Data Processing System (FDPS) also supplies warning tones that alert the crew to specific events or system failures. General Description Central Warning System Refer Figure - INDICATING AND RECORDING BLOCK DIAGRAM The central warning system is divided into the two parts that follow: The indicating and recording system has the sub−systems that follow:  Caution and Warning Lights  Clocks  Take−off Warning.  Flight Data Recorder System (FDR) The caution and warning light system is divided into the two parts that follow:  Central Computer  Central Warning System  Caution and warning lights  Electronic Instruments System.  Advisory lights. Caution and Warning Lights: The caution and warning lights system shows system malfunctions and other conditions that require a corrective action. REV 2.0 20 Sep 2024 PAGE 4 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Advisory Lights: The advisory lights show malfunctions and other conditions that require a corrective action and safe and normal system operation. The take−off warning system supplies an aural warning when a take−off is attempted with the aircraft not in the correct take−off configuration. Electronic Instruments System The Electronic Instrument System (EIS) shows navigation, engine, and system parameters. It interfaces with other systems to calculate, make, and show their images. The Electronic Instrument System (EIS) also monitors is calculations to prevent misleading information from being shown. REV 2.0 20 Sep 2024 PAGE 5 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - INDICATING AND RECORDING BLOCK DIAGRAM REV 2.0 20 Sep 2024 PAGE 6 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION ENGINE COCKPIT INTERFACE UNIT DHC-8-400 B1 & B2 TRAINING MANUAL ENGINE COCKPIT INTERFACE UNIT General Refer Figure - ENGINE COCKPIT INTERFACE UNIT The engine cockpit interface unit (ECIU) transfers data between the cockpit and both full authority digital engine controllers (FADECs). General Description Refer Figure - ENGINE COCKPIT INTERFACE UNIT (ECIU) LOCATION The ECIU is located in the flight compartment, on the lower shelf of the left circuit breaker console. It contains a single main circuit card assembly (CCA) that consists of two redundant channels (A and B). The ECIU receives 24 inputs per channel from the cockpit that are transmitted to the FADEC and 16 inputs per channel from the engines that are transmitted to the cockpit. Detailed Description Refer Figure - ENGINE COCKPIT INTERFACE UNIT − SYSTEM BLOCK DIAGRAM Inputs to the ECIU from the cockpit are hardware discretes and supply either a ground or an open indication. The ECIU monitors these discrete inputs and processes the information into ARINC data. The ECIU then transmits this data across four independent buses to each channel of both FADECs. The cockpit The ECIU can supply 16 discrete outputs per channel as either a ground or open discrete inputs are shown below in the table. indication. A 1.0−ampere discrete output drives the oil ejector solenoid while a 0.5−ampere output drives the other systems. The ECIU monitors the ARINC data and sets these discretes when the appropriate bits are set. The discrete outputs are shown in Table below. REV 2.0 20 Sep 2024 PAGE 7 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL breaker panel. T1.8 information is also input to the ECIU from each engine FADEC channel (four per aircraft) across the ARINC 429 data buses. The ECIU processes the T1.8 signal and supplies this information on the four ARINC data buses. The dual channel ECIU operates on 28 Vdc electrical power through 3−ampere circuit breakers. Channel A receives power from the left essential bus and channel B receives power from the right essential bus. The circuit breaker for channel A is located at position H5 on the left DC circuit breaker panel. The circuit breaker for channel B is located at position H5 on the right DC circuit REV 2.0 20 Sep 2024 PAGE 8 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - ENGINE COCKPIT INTERFACE UNIT REV 2.0 20 Sep 2024 PAGE 9 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - ENGINE COCKPIT INTERFACE UNIT (ECIU) LOCATION REV 2.0 20 Sep 2024 PAGE 10 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - ENGINE COCKPIT INTERFACE UNIT − SYSTEM BLOCK DIAGRAM REV 2.0 20 Sep 2024 PAGE 11 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION INDEPENDENT INSTRUMENTS DHC-8-400 B1 & B2 TRAINING MANUAL INDEPENDENT INSTRUMENTS General Description  LOC hours  Days Refer Figure - CLOCK SYSTEM BLOCK DIAGRAM  Months  Years (default on power−up is 90). The electronic clock supplies the following functions that follow: At each momentary activation of the ET switch, the applicable area of the  Continuous digital display of GMT or LOC display flashes and the data is then entered using the CHR button.  Digital display of ET  Display of the chronometer (CHR) function. In this mode, the display of The Elapsed Time (ET) switch located in the lower left area of the clock face seconds is analog (sweep−hand) and that of minutes is digital. gives 3−state and 2−state sequences dependent on the Weight on Wheels  Display of date and year, when requested. (WOW) status of the aircraft. There are two clocks located in the flight compartment one on each side of On the ground: the glare shield. The pilots set the type of time−based information to be shown on the display.  First activation: Display of Elapsed Time  Second Activation: Elapsed Time is reset to zero Detailed Description  Third activation: Display of Chronometer minutes Refer Figure - ELECTRONIC CLOCK Airborne: Each clock is independently operated from the switches located on the bezel.  First Activation: Display of Elapsed Time  Second Activation: Display of Chronometer minutes. The function selector positions are labeled DATE, LOC, GMT and SET. The selector is pushed to exit the SET mode. When the function selector is placed in the SET position, the Elapsed Time (ET) button is pushed to cycle through the modes that follow:  GMT minutes (displayed immediately when SET is selected)  GMT hours  LOC minutes REV 2.0 20 Sep 2024 PAGE 12 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL The Chronometer function switch (CHR) located in the top right corner of the When the selector switch is set to DATE, the day and month are shown in the clock face supplies the three states, in order, that follow: top 4−digit area of the clock face. The two left digits (01 to 12) identify the month and the two right digits identifies the day (01 to 31) and the year (0 to 99). As the day and year occupy the same area, the display alternates each second between the two parameters. To aid interpretation while displaying the year, the left digits are blank. Leap years are programmed into clock operation. Elapsed Time (ET) is indicated from 0 to 99:59 in the lower digital display area of the clock face and gives an indication of aircraft flight time. The mode is automatically enabled by a discrete Weight−Off−Wheels discrete input from the Proximity Switch Electronics Unit (PSEU) when the aircraft becomes airborne and can only be reset during a Weight−On−Wheels (WOW) When primary electrical power is removed, the time base is maintained by condition. A colon separates the hours and minutes. the aircraft battery bus, all displays are blanked, and the sweep−hand, if active, stops. Current parameters continue to increment with the exception Minutes are indicated from 0 to 59 by the two right digits in the lower display of the Chronometer and Elapsed Time functions. When primary power is area of the clock face with the left digits blanked. Seconds are shown against restored, the upper LCD display shows the original function data and the the round dial of the clock face by a sweep−hand activated by a stepper lower display indicates 00 00. The Chronometer sweep−hand returns to zero motor. and can be re−enabled if set to start from zero. The ARINC 429 output bus is not active during standby power operation. Greenwich Mean Time (GMT) is shown in the top 4−digit readout area of the clock face from 00:00 to 23:59 minutes when the function selector is in the GMT position. A single dot is displayed above the GMT legend. Local time (from 00:00 to 23:59) is shown in the same location as GMT when the selector switch is in the LOC position. A single dot appears above the LOC legend, to give an alternative means of distinguishing local time from GMT, additional to switch position. REV 2.0 20 Sep 2024 PAGE 13 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - CLOCK SYSTEM BLOCK DIAGRAM REV 2.0 20 Sep 2024 PAGE 14 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - ELECTRONIC CLOCK REV 2.0 20 Sep 2024 PAGE 15 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - Electronic Clock (Post ModSum 4-126403) Mode: The MODE pushbutton switch allows selection of the various modes in The clock has the indications that follow: turn:  UTC  GPS: In GPS mode the clock is synchronized with the GPS, if the GPS is  CHR valid. Time and date are updated.  ET.  INT: In the INT mode the clock runs in internal mode and uses its own time base if GPS is invalid. The default is to run in GPS mode. UTC: The UTC time is displayed from 0 to 23 hours, 59 minutes and 59 seconds  LT: In LT mode the local time is shown and the clock uses its internal on the six digit LCD (upper). time base.  DT: In this mode the date is shown and the clock uses the GPS time or The display can be: its own internal time base.  Synchronized with GPS if GPS is valid (GPS flag is shown) At start−up the clock automatically starts in the INT mode. If the GPS is valid  Internal time if GPS is invalid (INT flag is shown) the clock is automatically updated and switched in the GPS mode.  Date, if this mode is selected (DT flag shown) (date format is mm/dd/yy)  Local time, if this mode is selected (LT flag is shown). The time can be set with the SET function which is made available by pressing the MODE pushbutton for a minimum of two seconds. In this function the UTC minute digits flash (the INT flag is on). Use the ET RST pushbutton to increase The colon between the hours digits and the minutes digits is lit when the clock the minute indication; use the ET SEL pushbutton to decrease the minute is on. indication. While in this function, each push of the MODE button will cycle through: CHR: The chronometric time (minutes and seconds) is shown on the lower four−digit LCD from 0 to 99 hours, 59 seconds. The colon between the  The hour indication (the INT flag is on) minutes and seconds is lit when the chronometer is on. The display is blanked  The year indication (the DATE flag is on) when reset.  The month indication (the DATE flag is on)  The day indication (the DATE flag is on) ET: Elapsed time is displayed from 0 to 99 hours, 59 minutes on the lower  The LT minute indication (the LT flag is on) four−digit LCD. The elapsed time indication starts automatically when the  The LT hour indication (the LT flag is on). aircraft is weight off wheels. The colon is lit when the ET indicator is on. Reset of the display is possible only when the aircraft is weight on wheels. For each of these indications, use the ET RST pushbutton to increase the value and the ET SEL pushbutton to decrease the value. REV 2.0 20 Sep 2024 PAGE 16 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL CHR: This function operates with the CHR pushbutton switch which, with each The date is also supplied to the Flight Data Processing System (FDPS) located push, cycles through start (colon is lit), stop (colon is blanked) and reset in the Integrated Flight Cabinet (IFC1) to show when systems malfunctioned (display is blanked). If the display was occupied by the ET function before CHR for events in Central Diagnostics System (CDS) BITE reports. is selected, the indicator flag toggles between ET and CHR. The clocks send data to the CVR and FDPS through an ARINC 429 data bus. ET SEL and ET RST: The elapsed time function is operated with these two pushbutton switches. If the display was occupied by the CHR function before The clock installations interface with the aircraft systems that follow: ET SEL is selected, the indicator flag will toggle between CHR and ET. A push of the ET RST button blanks the ET display (only when the aircraft is weight on  IOP1, located in IFC 1, for input to the CDS and FDR wheels).  PSEU for Weight−Off−Wheels discrete (to activate the Elapsed Time function) The left essential and battery power busses supply 28 Vdc electrical power  CVR for recording of real time and synchronization with the FDR. through two 1 ampere circuit breakers to the CLOCK 1. The right main and battery power busses supply 28 Vdc electrical power through two 1 ampere circuit breakers to the CLOCK 2. Primary power is supplied from the left essential and right main 28 Vdc busses for CLOCK1 and CLOCK2 with a keep−alive voltage automatically supplied by the battery power bus when the primary electrical power is removed. The clocks are also supplied with variable 5 Vdc from the aircraft dimming bus to 4 display backlighting bulbs. Each clock operates independently. The CLOCK1 is interfaced directly with the Cockpit Voice Recorder and both clocks are interfaced with the Flight Data Recorder (FDR) through the Flight Data Processing System (FDPS). The FDR normally records time from the CLOCK1 but will switch to CLOCK2 if the CLOCK1 malfunctions. Real time is recorded on both the Cockpit Voice Recorder (CVR) and Flight Data Recorder (FDR) to establish synchronization between the two recording systems. REV 2.0 20 Sep 2024 PAGE 17 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - Electronic Clock (Post ModSum 4-126403) REV 2.0 20 Sep 2024 PAGE 18 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Clock, Electronic Refer Figure - ELECTRONIC CLOCK LOCATOR Two clocks are installed on the left and right sides of the glare shield panel assembly. On aircraft without ModSum 4−126403 and 4−126434 incorporated, the clocks have dichroic LCDs that show white digits against a grey background. On aircraft with ModSum 4−126403 or 4−126434 incorporated, the clocks have one six−digit and one−four digit LCDs that show white digits against a black background. REV 2.0 20 Sep 2024 PAGE 19 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - ELECTRONIC CLOCK LOCATOR REV 2.0 20 Sep 2024 PAGE 20 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION RECORDERS DHC-8-400 B1 & B2 TRAINING MANUAL RECORDERS Pushing a mechanical reset button on the switch assembly manually resets Introduction the switch following activation. The Solid State Flight Data Recorder (SSFDR) records aircraft parameter data The Flight Data Recorder System (FDR) has the components that follow: and stores it in crash−protected memory for future retrieval purposes.  Recorder, Flight Data General Description  Tray, Mounting  Switch, Inertia Figure - FLIGHT DATA RECORDER BLOCK DIAGRAM  Switch, Test Figure - Flight Data Recorder (Universal) Block Diagram  Switch, Anti-collision. The Solid State Flight Data Recorder (SSFDR) is a crash survivable unit that can An Underwater Locator Beacon (ULB) (also known as Underwater Locating record up to 25 hours of aircraft parameters and clock data. Device (ULD)) is attached to the Solid State Flight Data Recorder (SSFDR) handle. When immersed in either fresh water or saltwater, the ULB transmits The Solid State Flight Data Recorder (SSFDR) receives from the Integrated an acoustic signal to help locate the SSFDR. Flight Cabinet (IFC 1), the aircraft parameters that follow: Detailed Description  Flight path  Speed The flight data recorder system has SSFDR and an acceleration inertia switch.  Attitude The system interfaces with:  Engine power  Configuration  Flight Data Processing System (FDPS)  Operation.  Test circuit  Input power interlock arrangement An inertia switch removes power to the SSFDR and Solid−State Cockpit Voice  Caution and warning panel. Recorder (SSCVR) when exposed to a high acceleration. Pushing a mechanical reset button on the switch assembly manually resets the switch following The ULB is attached to the front of the SSFDR to easily access it for servicing activation. and to read its battery expiratory date. It may also be used as a carrying handle. On aircraft with ModSum 4Q126473 incorporated, two inertia switches are there, one each, for both the SSFDR and SSCVR is installed, which removes The ULB is automatically activated when immersed in either fresh water or power to the SSFDR and SSCVR when exposed to a high acceleration. salt water at depths from 0.5 to 20,000 feet (0.15 to 6096 meters). REV 2.0 20 Sep 2024 PAGE 21 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL When the ULB is activated, it will transmits an acoustic signal of 37.5 ±1 kHz The SSFDR operates in the modes that follow: frequency for a duration of 90 days or more. The ULB is self−powered by a battery with a typical service life of seven years from the date of manufacture.  Power off mode A label that shows the battery expiry date is attached to the ULB.  Initialization mode  Record mode The SSFDR system receives mandatory and non−mandatory parameters and  Continuous monitor mode discrete from aircraft systems at 128 words per second through an ARINC 717  Test mode data bus. The unit records up to 25 hours of data in a crash−survivable  Download mode memory module for subsequent retrieval and analysis. Ground−Based  Power down mode. Equipment (GBE) and an RS422 interface are used to download the recorded data. Power Off: The SSFDR operates when power is removed for more than 200 milliseconds. No functions are available. Applying power exits the power off On aircraft with ModSum 4Q126473 OR SB84−31−82 incorporated, the SSFDR mode. system receives mandatory and non−mandatory parameters and discrete from aircraft systems at 512 words per second through an ARINC 717 data Initialization: The SSFDR operates in the initialization mode immediately upon bus. Ground−Based Equipment (GBE) and an Ethernet interface are used to receiving power. It initializes Built−In−Test (BIT) within 500 millisecond (On download the recorded data. aircraft with ModSum 4Q126473 OR SB84−31−82 incorporated the initialization is completed within 250 millisecond). It determines the SSFDR The GBE performs the functions that follow: status.  Enables the real−time monitoring of aircraft parameters and discrete If the BIT fails, the external BIT light on the front panel of the SSFDR comes  Shows the internal status of the SSFDR on. The SSFDR exits the initialization mode of operation if it cannot record  Downloads data from the Crash Survivable Memory Unit (CSMU) data due to a critical failure. It transmits a signal to the continuous monitor  Does a functional test of the SSFDR. mode to stop it from recording, and the FLT DATA RECORDER annunciator on the Caution Warning Panel comes on. The Central Diagnostics System (CDS) receives and records an equivalent maintenance output through the The GBE interface connector (with protective cover) is installed in the front Integrated Flight Cabinet (IFC 1). panel of the SSFDR and is used for download, test and maintenance functions without removing the LRU from the aircraft. Record Mode: Recording data begins immediately following the initialization and self−test mode. The SSFDR stops recording if it does not receive data for more than five seconds, causing the FLT DATA RECORDER caution light to come on. REV 2.0 20 Sep 2024 PAGE 22 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL The SSFDR stays in the record mode until it is connected to the Ground Based Equipment (GBE). The type of GBE selects what mode the SSFDR will enter Down Load Mode: The SSFDR starts the down load mode when initiated by next. the external Down Load Unit (DLU). The SSFDR transmits the received data back to the Input/output Processor Power down Mode: The SSFDR power supply maintains internal voltages to (IOP1). This makes sure that the SSFDR is properly receiving the input data. If continue normal operation during a power interruption lasting less than 200 the status discrete senses a failure, the "loopback" data is not present. The milliseconds. For power interruptions greater than 200 milliseconds: FLT DATA RECORDER caution light on the caution and warning panel will come on.  The data in the RAM is lost  The initialization mode starts. The conditions that follow exit the SSFDR from the record mode: The SSFDR is electrically powered by 28 VDC through an interlock system. The  Removing power SSFDR starts to record when any of the conditions are met that follow:  Detection of the download mode  Command from the IOP.  Red anti-collision lights are switched ON  White anti-collision lights are switched ON Monitor Mode: The monitor mode supplies aircraft installation diagnostics  Both engines are operating, engine−operating oil pressure is sensed and troubleshooting. from both engines  Weight−Off−Wheels is sensed The SSFDR continuously does the following:  Test switch is set to GND TEST.  Background tests on system hardware  Monitors data inputs. In the monitor mode, the SSFDR records normal flight data in parallel with the data bus to the GBE. The GBE commands the SSFDR to operate in the monitor mode and to output a "data frame", once per second, that shows the status of the recorder, and the record data, 8 times per second. Test Mode: The test mode supplies "Return to Service" testing of the SSFDR. The external Automated Test Unit (ATU) commands the SSFDR to operate in the test mode. The normal record mode stops when the unit is operating in the test mode. The ATU controls the exit from the test mode. REV 2.0 20 Sep 2024 PAGE 23 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FLIGHT DATA RECORDER BLOCK DIAGRAM REV 2.0 20 Sep 2024 PAGE 24 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - Flight Data Recorder (Universal) Block Diagram REV 2.0 20 Sep 2024 PAGE 25 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - Flight Data Test Switch The FLIGHT DATA RCDR toggle switch on the Flight Data Recorder Panel is set to the GND TEST position to test the SSFDR while the aircraft is on the ground. Figure - Flight Data Test Switch REV 2.0 20 Sep 2024 PAGE 26 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - FDR CAUTION INDICATION If the SSFDR BITE senses satisfactory system operation, the FLT DATA RECORDER indication in the caution and warning panel goes out. Figure - FDR CAUTION INDICATION REV 2.0 20 Sep 2024 PAGE 27 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - ANTI−COLLISION LIGHT SWITCH The RED−OFF−WHITE A/COL toggle switch on the Exterior Lights Panel is set to the WHITE or RED position to power the SSFDR. The SSFDR system interfaces with the aircraft systems/busses that follow:  Left Main 28 VDC bus through 1 A circuit breaker F3  On aircraft with ModSum 4Q126473 OR SB84−31−82 incorporated, Right Essential 28 VDC bus through 5 A circuit breaker E10.  Avionics circuit breaker panel  IOM1 (located in IFC1) for input to the Centralized Diagnostic System (CDS) (maintenance flag)  IOP1 (located in IFC1) for input ARINC 573/717 data and loop−back data  Overhead anti-collision lights (red and white) switch for power interlock  Proximity Sensor Electronic Unit (PSEU) for WOW input (part of power interlock)  Overhead test switch  Caution panel (status flag)  Inertia switch (removes power for accelerations in excess of 5.5 g). REV 2.0 20 Sep 2024 PAGE 28 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - ANTI−COLLISION LIGHT SWITCH REV 2.0 20 Sep 2024 PAGE 29 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Recorder, Flight Data Refer Figure - FLIGHT DATA RECORDER LOCATOR Refer Figure - Flight Data Recorder (Universal) Locator The SSFDR has the components that follow:  Circuit Card Assemblies (CCAs)  Controller  Memory  Power supply regulator  Power supply filter. The SSFDR chassis is painted a bright international orange and is marked with the black letters "FLIGHT RECORDER DO NOT OPEN" and "ENREGISTREUR DE VOL NE PAS OUVRIR". Reflective tape is attached to the SSFDR external surface to further help recover it. The SSFDR is designed to operate in extreme environmental conditions with no forced air cooling. When installed, air can flow around the unit and supply cooling. The SSFDR is installed in the rear fuselage of the aircraft between station points X942.57 and X958.23 at Z134.46 and to the left of aircraft centerline. REV 2.0 20 Sep 2024 PAGE 30 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FLIGHT DATA RECORDER LOCATOR REV 2.0 20 Sep 2024 PAGE 31 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - Flight Data Recorder (Universal) Locator REV 2.0 20 Sep 2024 PAGE 32 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Tray, Mounting The SSFDR is installed in a ½ ATR short mounting tray that is electrically The current rating of each pole of the switch is 5 A at 28 VDC. bonded to the airframe by a bonding strap. NOTE On aircraft with ModSum 4Q126473 OR SB84−31−82 incorporated, the SSFDR tray is directly mounted on to the aircraft structure with help of four bolts, The maximum DC current of the SSFDR is less than 500 mA. which provides bonding to the chassis ground. The inertia switch connector is installed on the side panel of the SSFDR. The SSFDR mounting tray is located in the tail of the aircraft at stations X958.23 and Z134.43. The unit is installed at 45 degrees with the reset switch pointing towards ground and the front of the aircraft to sense upward and downward Switch, Inertia acceleration. It is attached to a bracket located below the cabin floor at station point X195.8, Y0, Z88.9. Refer Figure - Flight Data Recorder Inertia Switch Locator The inertia switch supplies the SSFDR system with the capability to remove power from the system if an acceleration of greater than 5.5 g is applied to the switch as a consequence of high impact landing. This makes sure that SSFDR recorded data prior to the impact is stored and cannot be overwritten. The normally−closed inertia switch has an impact activated double−pole latching switch; one pole of which is connected in series with the 28 VDC aircraft supply to the SSFDR. (The second pole is connected to the SSCVR). The switch latches to an open state when activated. On aircraft with ModSum 4Q126473 incorporated, two inertia switches are there, one each, for both the SSFDR and SSCVR is installed. The existing inertia switch provides 28 VDC power to the SSFDR and the new inertia switch, installed above the SSFDR inertia switch, provides the 28 VDC power to the SSCVR. A manually operated RESET BUTTON, located on the side of the inertia switch is used to re−arm the switch. REV 2.0 20 Sep 2024 PAGE 33 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - Flight Data Recorder Inertia Switch Locator REV 2.0 20 Sep 2024 PAGE 34 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Switch, Test Refer Figure - Flight Data Recorder Test Switch When the FLIGHT DATA RCDR switch is set to the GND TEST position, the normal "on ground" power removal circuits are bypassed. The pre−flight GND TEST switch located in the flight compartment on the overhead panel, supplies power to the SSFDR for the duration that the switch is held. If the SSFDR BITE indicates satisfactory system operation, the caution panel FLT DATA RECORDER light goes out. The NORM−GND TEST toggle switch is located in the flight compartment on the overhead FLIGHT DATA RCDR panel. Switch, Anti−Collision Refer Figure - ANTI−COLLISION LIGHT SWITCH When the anti-collision lights switch is set to the RED or WHITE position, the normal "on ground" power removal circuits are bypassed. The anti-collision lights switch is located in the flight compartment on the overhead panel. If the SSFDR BITE indicates satisfactory system operation, the caution panel FLT DATA RECORDER light goes out. The RED−OFF−WHITE A/COL switch is located in the flight compartment on the overhead EXTERIOR LIGHTS panel. REV 2.0 20 Sep 2024 PAGE 35 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - Flight Data Recorder Test Switch REV 2.0 20 Sep 2024 PAGE 36 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION SIGNAL CONDITIONING UNIT DHC-8-400 B1 & B2 TRAINING MANUAL SIGNAL CONDITIONING UNIT The FSCU supplies analog inputs to interface with the devices that follow: Introduction  Potentiometers  Linear Variable Differential Transformers (LVDTs) Refer Figure 31- 18 Signal Conditioning Unit  Pressure Sensors The discrete, force, position and pressure sensors measure the control The FSCU performs the tests to detect the critical faults that follow: column inputs and also the affecter response. The aircraft system sensors convert the control column inputs into electrical signals and transmit them to  Power supply fault Flight Data Recorder Signal Conditioning Unit (FSCU). The FSCU supplies  A/D converter fault excitation, demodulation and filtering signal for the sensing devices. The FSCU  SRAM fault also supplies ARINC 429 data for the Flight Data Recorder (FDR).  Program memory fault  EEPROM checksum fault General Description  Watchdog fault Refer Figure - Signal Conditioning Unit Block Diagram The FSCU operates in the states that follow:  On  Off The FSCU interfaces with the aircraft 28 VDC power supply. The power supply converts the input into the conditioned supply and reference voltages. The FSCU interfaces with the 24 discrete inputs and 4 discrete outputs. The FSCU receives data from at least five ARINC 429 receivers at 100 kbps. REV 2.0 20 Sep 2024 PAGE 37 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - Signal FigureConditioning Unit Block Diagram 31- 1 Signal Conditioning Unit REV 2.0 20 Sep 2024 PAGE 38 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION CENTRAL COMPUTER DHC-8-400 B1 & B2 TRAINING MANUAL CENTRAL COMPUTER Each Flight Data Processing System (FDPS 1, FDPS 2) has the modules that Introduction follow: The central computer has a Flight Data Processing System (FDPS) to receive  Integrated Flight Cabinet data from sensors and avionics systems and supply it to other systems. The  Input/output Processor Flight Data Processing System (FDPS) also causes different warning tones to  Input/output Module sound if an important system has malfunctioned or if the aircraft is in a  Prime Power Supply Module dangerous condition.  Aircraft Configuration Module. General Description The modules are located in two Integrated Flight Cabinets (IFC 1, IFC 2) installed in the Avionics rack. Refer Figure - FLIGHT DATA PROCESS SYSTEM BLOCK DIAGRAM Refer Figure - FLIGHT DATA PROCESS SYSTEM BLOCK DIAGRAM, POWER The two Flight Data Processing System (FDPS1, FDPS2) do the functions that follow:  Receives and calculates aircraft and avionics parameters for the Flight Data Recorder (FDR)  Receives data from different avionics and aircraft systems and routes their parameters to another systems through a single output  Receives analogue, discrete, and digital data from other systems and changes it to ARINC 429 format  Mismatch message calculations  Monitors the Electronic Instrument System (EIS)  Message generation  Aircraft Configuration Management (ACM)  Supplies Warning Tones (WTG 1, WTG 2) and manages its priority  Maintenance tests  Software teleloading. REV 2.0 20 Sep 2024 PAGE 39 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FLIGHT DATA PROCESS SYSTEM BLOCK DIAGRAM REV 2.0 20 Sep 2024 PAGE 40 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FLIGHT DATA PROCESS SYSTEM BLOCK DIAGRAM, POWER REV 2.0 20 Sep 2024 PAGE 41 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Detailed Description The POST checks the parameters that follows: The Flight Data Processing System (FDPS 1, FDPS 2) functions in the modes that follow:  Program memory  Data memory  Initialization (INIT)  Watch dog  Power−On Self-Test (POST)  Power supplies Hardware (HW) monitoring  Operational (OPER)  Module Validity HW and Software (SW) logic  Maintenance  Memory partitioning HW monitoring  Teleloading.  Time partitioning HW monitoring. Initialization (INIT) State: The FDPS functions in the initialization mode after If a parameter malfunctions, it will causes the FDPS to malfunction. an electrical power interruption. It initializes the hardware and software when it is on the ground. The POST also checks the parameters that follow: If the FDPS does not receive a maintenance or teleloading request, it will then  ARINC 429 inputs and outputs function in the Power−On Self-Test (POST) mode.  ARINC 422 inputs and outputs Power−On Self-Test (POST): The FDPS does a (POST) to make sure the  ARINC 717 inputs and outputs hardware operates correctly before starting the operational mode.  Discrete inputs and outputs  IO1, IO2 board validity The FDPS does a POST when the aircraft is on the ground and there is a long  Analogue inputs and outputs electrical power interruption that continues for more than 200 milliseconds.  Discrete inputs and outputs. The FDPS tests the interfaces that follow: They do not cause a FDPS malfunction.  Input/output Processor (IOP 1, IOP 2) The POST sequence continues for 25 seconds.  Input/output Module (IOM 1, IOM 2)  Ground Proximity Warning System Converter. The result of the POST is stored in the Built In Test Equipment (BITE) and sent to the Central Diagnostic System (CDS). REV 2.0 20 Sep 2024 PAGE 42 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Operational State: The FDPS does the functions that follow: Two specific external non−avionics General Purpose Data Buses (GPDB 1, GPDB 2) are used to transmit data to systems that are not part of the avionics  Flight Data Concentrator (FDC) suite.  Data Hub Concentrator (DHC)  Data Control (DCO) Data Control (DCO): The Data Control (DCO) makes calculations for the  Mismatch calculations parameters before concentration.  Electronic Instrument System (EIS) essential monitoring  Advisory message generation  Aircraft configuration management  Warning Tone Generator (WTG 1, WTG 2). Flight Data Concentrator (FDC): The FDC is located in IOP1. It receives data in different formats from the avionics and other aircraft systems and supplies digital data to the Solid State Flight Data Recorder (SSFDR) through an ARINC 573/717 bus. Data Hub Concentrator (DHC): The FDPS receives noncritical data from avionics and other aircraft systems and supplies their parameters to other systems through a single output. The FDPS supplies the concentrated parameters through ARINC 429 data buses to the avionics systems that follow:  Electronic Instruments System (EIS)  Flight Guidance Module (FGM 1, FGM 2)  Stall Protection Module (SPM 1, SPM 2)  Audio and Radio Control Display Units (ARCDU 1, ARCDU 2)  Traffic Collision Avoidance System (TCAS)  Ground Proximity Warning System (GPWS)  Weather Radar (WXR)  Flight Management System (FMS 1, FMS 2). REV 2.0 20 Sep 2024 PAGE 43 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - EIS ESSENTIAL MONITORING REV 2.0 20 Sep 2024 PAGE 44 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL to show a malfunction condition for that parameter. Refer Figure - EIS ESSENTIAL MONITORING Refer Figure - FDPS MESSAGE GENERATION Mismatch Calculations: The FDPS monitors different parameters. If an IOP senses a difference between its related value and the same the parameters it The essential monitoring parameter thresholds are shown in the table that received from the other IOP, a mismatch message is shown by the Flight Mode follows: Annunciators (FMA) in the opposite Primary Flight Displays (PFD). The Flight Data Processing System (FDPS) Mismatch Messages that follow:  PITCH MISMATCH  HEADING MISMATCH  IAS MISMATCH  ALT MISMATCH  RAD ALT MISMATCH  GS MISMATCH  LOC MISMATCH. The calculations are done only when the parameters are valid. For localizer Message Generation: The Message Generation System has the modes that and glideslope mismatch messages, the two navigation receivers outputs follow: must be valid and tuned to the same frequency.  Advisory  Flight Mode Annunciator (FMA). Electronic Instrument System (EIS) essential monitoring: The FDPS does essential monitoring for the systems that follow: Advisory: The advisory messages are shown in white letters on the Engine Display (ED). Each message has a location and is shown while the condition is  Radio Altimeter present. There are two kinds of advisory messages as follows:  Localizer deviation  Glideslope deviation.  First Family  Second Family. The IOPs monitors the difference between a parameter received directly and same parameter received from the opposite PFD and IOP. When the First Family: The first family messages relate to an applicable aircraft status, difference is more than the predetermined value, the FDPS causes to the PFD where crew awareness is required and an action may be necessary. These REV 2.0 20 Sep 2024 PAGE 45 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL messages are located near its related indication. Second Family: Second family messages relate to minor malfunctions and are IFC Messages: The most important IFC message is shown at the bottom left part located at the bottom of the ED. The messages relate to passive malfunctions of the ED. It shows the messages that follow: do not affect a continued safe flight.  IOP1 FAIL, IOP2 FAIL, IOPS FAIL The FDPS causes the ED to show the advisory messages that follow:  IOP BAD CONF  IOM1 FAIL, IOM2 FAIL, IOMS FAIL  [BALANCE]  WTG1 FAIL, WTG2 FAIL, WTGS FAIL  INCR REF SPEEDS  GPWSC I/F FAIL  ICE DETECTED  WOW/IOP1 FAIL, WOW/IOP2 FAIL, WOW/IOPS FAIL  IFC messages  IFC ACM1 FAIL, IFC ACM2 FAIL, IFC ACMS FAIL  Display messages.  RA1 FAIL, RA2 FAIL, RAS FAIL. NOTE IOP1 FAIL, IOP2 FAIL, IOPS FAIL: The IOP FAIL messages are shown when the IOP1 or IOP2 or the interface to the ED has malfunctioned. Note: The advisory messages are shown in order of decreasing importance. The most important IFC message is shown at the bottom left part of the ED IOP BAD CONF: The IOP BAD CONF message is shown when a bad aircraft and the most important display message is shown at the bottom right part configuration condition is sensed by one IOP or the other. The message comes of the ED. into view when the aircraft is on the ground. [BALANCE] Message: The BALANCE message comes into view flashing for five IOM1 FAIL, IOM2 FAIL, IOMS FAIL: The IOM FAIL messages are shown when IOM seconds then stays on steady when a fuel imbalance is sensed by the left or malfunctions are sensed. right fuel gauging computers. WTG1 FAIL, WTG2 FAIL, WTGS FAIL: The WTG FAIL messages are shown when INCR REF SPEEDS Message: The INCR REF SPEEDS message comes into view the WTG malfunctions. The message is shown when the aircraft is on the ground when the Stall Protection System (SPS) stall sensing is changed for icing and engines are not running. conditions. The FDPS sends a discrete data signal to Electronic Instruments System (EIS) from one or the other SPM. GPWSC I/F FAIL: The GPWSC I/F FAIL messages are shown when the IFC1 makes the GPWS inoperative when it malfunctions. The message is when the ICE DETECTED Message: The ICE DETECTED message comes into view when one malfunction is sensed. ice detector probe or the other senses ice accumulation. REV 2.0 20 Sep 2024 PAGE 46 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL WOW/IOP1 FAIL, WOW/IOP2 FAIL, WOW/IOPS FAIL: The WOW/IOP FAIL messages are shown when the IOP senses a difference between the main and nose WOW signals from the PSEU. The IOP will not be able to do a POST after a power interruption. The message is shown when the aircraft is on the ground and engines are not running. IFC ACM1 FAIL, IFC ACM2 FAIL, IFC ACMS FAIL: The IFC ACM FAIL messages are shown when a ACM malfunction is detected. A single ACM malfunction has no effect on the aircraft. When the two ACM malfunction, a bad DU BAD CONF messages is shown. RA1 FAIL, RA2 FAIL, RAS FAIL: The RA FAIL message is shown when a dual Radar Altimeter system is installed and the RA malfunction is detected for more than ten seconds by the EIS ED. The message is displayed any time the malfunction is detected. Display Messages: The most important display message is shown at the bottom right part of the ED. The FANS FAIL message is the only FDPS related display message indication. It comes into view when 2 or more avionics cooling fans do not operate. The FANS FAIL display message also comes into view when 2 or more avionics cooling fans do not operate and they are not inhibited by their related thermal switch while the aircraft is on the ground. The temperature switch 1 supplies data to IOP1 and temperature switch 2 supplies data to IOP2. An IFC or display advisory message also causes the AVIONICS caution light to come on 2 minutes after the aircraft is on the ground and the air speed is less than 50 kts. REV 2.0 20 Sep 2024 PAGE 47 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FDPS MESSAGE GENERATION REV 2.0 20 Sep 2024 PAGE 48 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FDPS AIRCRAFT CONFIGURATION MODULES REV 2.0 20 Sep 2024 PAGE 49 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - FDPS AIRCRAFT CONFIGURATION MODULES The tone that sounds depends on its priority. The most important tone always Aircraft Configuration Management: The FDPS uses data that is stored in the sound as follows: Aircraft Configuration Modules (ACM 1, ACM 2). The ACMs are attached to the back−panel of its related IFC1. Each IOP supplies the configuration data to the systems that follow:  Electronic Instrument System (EIS)  Stall Protection System (SPS), Stall Protection Modules (SPM 1, SPM 2)  Auto Flight Control System (AFCS), Flight Guidance Modules (FGM 1, FGM 2)  Audio and Radio Control Display Units (ARCDU 1, ARCDU 2) Each system monitors the configuration data. If a malfunction is sensed, it is stored in the Built in Test Equipment (BITE) and sent to the Central Diagnostics System (CDS) and the ED shows an IFC message. The Aircraft Configuration Modules (ACM 1, ACM 2) are programmed using the Portable Multipurpose Access Terminal (PMAT). Refer Figure - FDPS WARNING TONE GENERATORS (WTG1, WTG2) Warning Tone Generators (WTG 1, WTG 2): The Warning Tone Generators (WTG 1, WTG 2) give different tones to tell the pilots of dangerous conditions or system malfunctions. REV 2.0 20 Sep 2024 PAGE 50 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL The WTG is not an independent system. It is part of the Input/output Modules’ (IOM1, IOM2) function. One WTG supplies a warning tone to the Remote Control Audio Unit (RCAU) and other monitors it. The RCAU amplifies the signal and sends the tone to the flight compartment speakers and the pilots headsets. GPWS: The GPWS is a system that makes its own synthesized voice sounds and connects directly to the Audio Integrating System (AIS). When the WTG senses a GPWS audio on condition, it inhibits the other tones. The voice sounds from the GPWS is inhibited if a GPWS malfunction is sensed while the aircraft is airborne. The GPWS audio on signal is monitored to prevent an inhibit of a different WTG tone caused by a partial GPWS malfunction. If the GPWS audio on signal stays on for more than 60 seconds, the GPWS’s priority status is ignored. A GPWS and a TCAS or WTG tone is allowed to be heard at the same time. The GPWS malfunction condition and that the TCAS and WTG continues to function is easily identified by the pilots. To be able to sense this malfunction condition, the FDPS does not inhibit the GPWS during the WTG test (ADC1 or ADC2 TEST toggle switch selection). TCAS: The TCAS is a system that makes its own synthesized voice sounds and connects directly to the Audio Integrating System (AIS). It has two types of alerts with two different priority levels that follows: Note  Resolution advisories The WTGs generate the aural tones and control the priority of the voice  Traffic advisories. sound from the GPWS and TCAS. REV 2.0 20 Sep 2024 PAGE 51 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FDPS WARNING TONE GENERATORS (WTG1, WTG2) REV 2.0 20 Sep 2024 PAGE 52 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL The TCAS malfunction condition and that the GPWS and WTG continues to The WTG sounds a clicking tone when the pitch trim is in motion for more than function is easily identified by the pilots. To be able to sense this malfunction 3 seconds. condition, the FDPS does not inhibit the TCAS during the WTG test (ADC1 or ADC2 TEST toggle switch selection). The two Flight Control Electronic Control Units (FC ECU1, FC ECU2) supply pitch trim in motion data through data buses to the WTGs. Engine Fire: The WTG inhibits the engine fire aural warnings when a more important tone sounds. The clicking tone sounds if one FCS ECU or the other supplies a pitch trim in motion signal to the WTG. A continuous chime sounds when the WTG senses a fire bell discrete signal from the fire detection system. Overspeed Warning: The WTG inhibits the overspeed warning tones when a more important tone sounds. Figure - FDPS Incorrect Take−Off Configuration Warning The WTG sounds an intermittent 1000 Hz tone when the aircraft’s speed is more Incorrect Take−Off Warning: The WTG inhibits the incorrect take−off warning than Maximum Velocity in Operation (VMO). indication when more a more important tone sounds. The two Air Data Units (ADU1, ADU2) supply overspeed data through data buses Autopilot Disengagement: The WTG inhibits the autopilot disengagement tones to the WTGs. when a more important tone sounds. The overspeed tone sounds if one ADU or the other supplies an overspeed The WTG sounds an intermittent tone for 1.5 seconds for a manual AFCS signal to the WTG. disengagement. It sounds a 4000 Hz Intermittent tone when the Auto Flight Control System (AFCS) automatically disconnects until it is cancelled. The two Flight Guidance Modules (FGM1, FGM2) supply Autopilot (AP) engagement/disengagement data through data buses to the WTGs. The autopilot disengagement tones sound if one FGM or the other supplies a disengagement signal to the WTG. Pitch Trim in Motion: The WTG inhibits the pitch trim in motion tones when a more important tone sounds. REV 2.0 20 Sep 2024 PAGE 53 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FDPS Incorrect Take−Off Configuration Warning REV 2.0 20 Sep 2024 PAGE 54 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - FDPS INCORRECT LANDING GEAR CONFIGURATION Incorrect Landing Gear Configuration: The WTG inhibits the incorrect landing gear configuration tones when a more important tone sounds. The WTG sounds a continuous 800 Hz tone when the landing gear is not set to a safe for landing configuration. REV 2.0 20 Sep 2024 PAGE 55 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FDPS INCORRECT LANDING GEAR CONFIGURATION REV 2.0 20 Sep 2024 PAGE 56 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - FDPS ALTITUDE ALERT Altitude Alert: The WTG inhibits the altitude alert tones when a more important tone sounds. The WTG sounds a continuous 2900 Hz tone for 1 second when the aircraft enters an area that is less than 1000 ft. (305 m) above or below a set altitude. REV 2.0 20 Sep 2024 PAGE 57 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Figure - FDPS ALTITUDE ALERT REV 2.0 20 Sep 2024 PAGE 58 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL Refer Figure - FDPS BETA LOCKOUT WARNING Each WTG monitors its output. If the WTGs calculate different tones, the system will use the calculations from WTG 1. Its calculations are more important than Beta lockout warning: The WTG inhibits the beta lockout alerts when a more WTG 2. important tone sounds. The WTGs does the monitoring that follows: The Warning Tone Generator (WTG 1, WTG 2) sounds a continuous 800 Hz tone when the Power Lever Angle (PLA) is set below the IDLE position while  Software in flight. To give a beta lockout indication, the Flight Data Processing System  Hardware. (FDPS) senses the parameters that follow: Software Monitor: Each Warning Tone Generator (WTG 1, WTG 2) uses their  Beta lockout switch position related inputs to calculate the tone logic. The tone calculations are compared  Power lever angle and the WTG 1 supplies the applicable tone.  Main landing gear WOW. Hardware Monitor: Each Input/Output Module (IOM 1, IOM 2) has two Master Warning: The WTG inhibits the master warning tones when a more Input/Output boards (IO1, IO2). The WTG 1 sends an analogue tone to the important tone sounds. The WTG sounds three chimes when one or the other Remote Control Audio Unit (RCAU) through IO2. It routes the same signal to IO1 red master warning light comes on. to make sure that the output tone is correct. If there is a difference, the WTG 1 stops functioning and makes itself invalid. Then, the WTG 2 functions when Master Caution: The WTG inhibits the master caution tones when a more required. important tone sounds. The WTG sounds a single chime when one or the other amber master caution light comes on. The Warning Tone Generator (WTG 1, WTG 2) will remain off until the next power interruption when it is latched off because of a malfunction. The SELCAL: The WTG inhibits the SELCAL tone when a more important tone Power−On Self-Test (POST) will determine its validity. sounds. A continuous 1200 Hz tone sound for three seconds when the Selective Calling (SELCAL) system senses an incoming call. A Warning Tone Generator (WTG 1, WTG 2) malfunction is stored in the Built in Test Equipment (BITE) and sent to the Central Diagnostic System. The Engine There Warning Tone Generators (WTG 1, WTG 2). The WTG1 sounds the Display (ED) advisory message location will show a WTG FAIL message when a applicable tone when necessary while WTG2 functions in the standby mode. WTG malfunctions. The WTG2 functions only when it senses that WTG1 has malfunctioned. FDPS Abnormal Modes: Most aircraft and avionics system supplies data to other The WTG receives inputs from other systems to make it operate. systems through the two FDPS’s. Each aircraft or avionics system uses data from its related FDPS. REV 2.0 20 Sep 2024 PAGE 59 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL If its related FDPS malfunctions, the systems will automatically receive data Maintenance Mode: The Built in Test Equipment (BITE) uses the Central from the other FDPS. Diagnosis System (CDS) to give the condition of the component. It stores faults in a Non-Volatile Memory (NVM) for reporting to line and shop The Flight Data Recorder (FDR) receives data through the Input/Output maintenance. Processor (IOP 1) from the two FDPS’s. The Built in Test Equipment (BITE) allows aircraft maintenance personnel to: If FDPS1 malfunctions, the Flight Data Recorder (FDR) function also malfunctions. If FDPS 2 malfunctions, the FDR will not record its parameters.  Do fault isolation and return to service testing after completing maintenance actions The data is supplied only to the FDPS1 from the systems that follow:  Access failure reports from last or previous flight legs  Get the avionics status report  Hydraulic quantity 1  Get the part number of a given part  Hydraulic quantity 3  Fuel inlet temperature 1 The Built in Test Equipment (BITE) modes monitors the condition of the  Parking brake pressure component as follows:  Main Oil Pressure 1  Ground Proximity Warning System Converter.  Power−On Self-Test (POST)  Continuous Monitoring. If FDPS1 malfunctions, its parameters are shown as dashes on the Engine Display (ED). Power−On Self-Test (POST): The Power−On Self-Test (POST) checks the condition of the component at Power−Up or after a long power interruption. The data is supplied only to the FDPS2 from the systems that follow: Continuous Monitoring: The Continuous Monitoring checks the status of the  Hydraulic quantity 2 component in flight. It records faults in a Non-Volatile Memory (NVM) for  Fuel inlet temperature 2 later troubleshooting using the Central Diagnosis System (CDS). A long power  Traffic Collision Avoidance System (TCAS). interruption causes the FDPS to start a Power−On Self-Test (POST) again. If FDPS2 malfunctions, its parameters are shown as dashes on the Engine Teleloading: The CDS is a communication connection between the FDPS and Display (ED). a Portable Multipurpose Access Terminal (PMAT). The PMAT connects to a Personal Computer (PC) to download a new software version when a software If one FDPS or the other malfunctions, the Engine Display (ED) will also show an upgrade is necessary. IFC message in the advisory message area. REV 2.0 20 Sep 2024 PAGE 60 DGM12CBXX3Q4PW TRAINING PURPOSE ONLY  MAINTENANCE TRAINING ORGANISATION DHC-8-400 B1 & B2 TRAINING MANUAL The CDS functions in the teleloading mode when the conditions that follow is correct:  The Calibrated Air Speed (CAS) is less than 50 kts for more than 10 seconds  The aircraft is on the ground  The CDS GND MAINT toggle switch on maintenance panel is set  The MAINT Key on the ARCDU is pushed. The left essential bus supplies electrical power through a 10 A circuit breaker and the Prime Power Supply Module (PPSM1) to the Input/Output Processor (IOP1) and Input/Output Module (IOM1). The circuit breaker is located in position F7 on the avionics circuit breaker panel. The left main bus supplies electrical power through a 7.5 A circuit breaker and the PPSM1 to the Stall Protection Module (SPM1). The circuit breaker is located in position F2 on the avionics circuit breaker panel. The right main bus su