Navigation.docx

Navigation 2A-34--10: Navigation Systems General Description: The navigation systems provide the flight aew with indications of position in three dimensions and vector information for management of the flight environment, supplemented with aural and visual alerts to prevent Controlled Flight Into...

Navigation 2A-34--10: Navigation Systems General Description: The navigation systems provide the flight aew with indications of position in three dimensions and vector information for management of the flight environment, supplemented with aural and visual alerts to prevent Controlled Flight Into Terrain (CFiT) and collision with other airborne traffic. Flight environmental data, aircraft attitude and direction are integrated with the PlaneView display system, Automatic Flight Control System (AFCS) and the Multi-Function Control Display Units (MCDUs). Full discussions of display formats, autopilot I autothrottle and tuning procedures are contained in sections 2B-02-00, 28-04-00, 28-05-00 and 28-06-00 of this manual.This section details the sensor systems used to determine airspeed and altitude, standby instrumentation providing backup indications of airspeed, altitude, attitude and heading, the integration of the Air Data System (ADS) that supplies data to the AFCS and the onboard systems that provide alerts and warnings to prevent hazardous fight conflicts. The Enhanced Ground Proximity Warning System (EGPWS) is a terrain awareness and warning system incorporating alerting and display functions. The system uses aircraft geographic position, altitude, climb and descent rate, and a terraindatabase to determine potential conflicts between the aircraft flight path and terrain, and to provide aural alerts and visual depictions of hazardous terrain dearance. The Traffic and Collision Avoidance System (TCAS) uses transponder signal information to detect and plot the tracks of other airborne traffic and formulate flight guidance for maneuvers to avoid potential conflicts. The Visual Guidance System or Head Up Display (HUD) allows the pilot to maintainvisual contact with the external environment while maintaining a full instrument scan by superimposing instrument data onto a transparent panel placed at the pilot eye level. The Enhanced Vision System (EVS) amplifies infra-red light emanating from ground sources to provide the pilot with greater situational awareness during periods of darkness or low visibility. The navigation system is divided into the following subsystems: 2A-34-20: Flight Environment Data System 2A-34-30: Standby Multifunction Controller (SMC) System 2A-34-40: Head Up Display (HUD) II System 2A-34-50: Enhanced Vision System (EVS) II Crew Alerting System (CAS) Messages: Although the integrated navigational systems are discussed in Sections 2B-02-00, 2B-04-00, 28-05-00 and 28-06-00 of this manual, the following tabulation of CAS messages relating to the performance of navigational systems and radios is provided for the convenience of the reader: Area Monitored CAS Message Navigation Systems Comparison @1fjle'f @ (caution) Inertial Reference Systems Inertial Reference Systems llitjffljjjj (caution) Inertial Reference Systems IRS 1-2-3 Pos Ent (caution) Lateral Navigation Lateral Cpl Data Invalid (caution) Vertical Navigation Vertical Cpl Data lnvali (caution) ADF Systems 'i1tJiiD (advisory) Navigation Systems Comparison CAT 2 AGM Reversio (advisory) Navigation Systems I RadarAltimeter Comparison [lfJfj•J: (advisory) Navigation Systems I RadarAltimeter Comparison (+f!jfJ;tGt!ll(advisory) Navigation Systems Comparison [if!jfj;Qlt:fl (advisory) Distance Measuring Equipment i•@lfjitjl (advisory) Flight Guidance Computer FGC 1-2 Maste (advisory) Flight Guidance Computer IAir Data System Interface FGC 1 Not Using ADS 1-2· (advisory) Flight Guidance Computer IAir Data System Interface FGC 2 Not Using ADS 1-2- (advisory) Flight Guidance Computer / IRS Interface FGC 1 Not Using IRS 1-2- (advisory) Flight Guidance Computer / IRS Interface FGC 2 Not Using IRS 1-2- (advisory) Flight Guidance Computer I Sensor I Radio Interface FGC-NAV Miscompa (advisory) Flight Guidance Computer I RadAlt Interface FGC - AfT Rad Alt lnvali (advisory) Flight Management System l@.'i!•jfflitjl (advisory) Global Positioning Satellite System @fjfjCI {advisory) AGM 1-2-3-4 DMU Charts Fai (advisory) Navigational Databases Inertial Reference Systems IRS 1-2-3 Alignin (advisory) Inertial Reference Systems lltJj•QICI (advisory) Inertial Reference Systems IRS 1-2-3 Pos Ent (advisory) Area Monitored CAS Message Inertial Reference Systems !i;JlletjfM@ {advisory) Inertial Reference Systems IRU Sec Pwr 1-2-3 Fai (advisory) VHF Navigation Radios !m)*jfJttJ! (advisory} Vertical Navigation Path VNAV Track Chang (advisory) * Numerous CAS messages address rurrency and acruracy of currently loaded navigational databases in the Flight Management System (FMS). The text of the message, represented by XXXXX, indicates the nature of the invalid data in the databases. 2A-34-20: Flight Environment Data System General Description: (See Figure 1.Air Data System (ADS) Block Diagram.) The flight environment data system senses environmental conditions to determine and display aircraft attitude and direction. The system includes the following components: Digital Air Data Computing Static Air Temperature I Total Air Temperature Sensing Total Air Temperature Probe I Cabin Pressure Relief Air Valve Description of Subsystems, Units and Components: Digital Air Data Computing (DADC): The DADC system measures the pressure and temperature of the external air and supplies this data to the other systems. It consists of four Multi Function Probes I Air Data Computers (MFP I ADCs) and two dual element Total Air Temperature (TAT) probes. The DADC system provides accurate altitude, calibrated airspeed, vertical speed, true airspeed, mach number, angle of attack, angle of side slip, air pressure and air temperature data essential for the control, navigation and propulsion of the aircraft. DADC system architecture eliminates the need for separate Angle of Attack (AOA) sensors, static ports, air data computers, current regulators and current monitors typical of a traditional system. In addition, pneumatic plumbing to remote sensors and the associated fittings and drains are eliminated. The DADC system consists of four integrated Multi-function Probes (MFPs) and two dual element Total Air Temperature (TAT) probes. The DADC system provides accurate altitude, calibrated airspeed, vertical speed, true airspeed, mach number, AOA, angle of side slip, air pressure and air temperature data essential for the control, navigation and propulsion of the aircraft. The system measures the pressure and temperature of the external air and supplies this data to the other systems. Multi-function Probe: Each MFP has two independent Air Data (AD) channels within it, an AD channel which calculates onside air data and local AOA information and an Opposite Side Pressure (OSP) channel which provides static pressure to the AD channel on the other side of the aircraft. The primary functions of the AD channels are as follows: Measures local pressures Receives TAT information from the TAT probes Receives opposite side pressure for beta compensation and angle of side slip calculation Calculates aircraft air data parameters and true AOA Calculates the Static Source Error Correction (SSEC) for the aircraft AD parameters Transmits results and status to the appropriate aircraft systems (via ARINC 429 interface) Receives aircraft status and configuration (via ARINC 429 interface) Shares control and monitors the MFP heater current The primary functions of the OSP channel are as follows: Measures local pressures Transmits results and status to the AD channel on the opposite side of the aircraft Shares control and monitors the MFP heater current The MFP contains the pneumatic porting necessary to deliver impact, static and AOA pressures to the AD channel pressure sensors. The MFP has a single integral deicing I anti-icing heater element. The AD channel and the OSP channel have independent static ports and independent transducers. The MFP measures the current and voltage of the heater power and regulates the on time of the heater to maintain a probe tip temperature of 150°C ±50°C when airspeed is less than 60 knots. Above 60 knots, the MFP heater is turned full on. Total Air Temperature Probe: Two dual element TAT probes located on the left and right sides of the forward fuselage provide outside air temperature to the associated MFP. The MFP controls and monitors the TAT heater to reduce the power requirements based on temperature and aircraft state. Total Air Temperature Probe I Cabin Pressure Relief Valve: The total air temperature probe I cabin pressure relief valve is activated when the WOW relay is in the ground mode. The valve provides bleed air to aspirate the TAT probes during ground operation. The valve also routes warm bleed air to open the cabin pressure relief valve during ground operation. Static Air Temperature (SAT)/ Total Air Temperature (TAT) Sensing: The TAT probes provide the MFP I ADCs with ambient temperature data. Static airtemperature is calculated by the air data function within the MFP I ADCs. The MFP I ADCs process this temperature data and disperses this information to the various systems and instruments requiring this data. The TAT sensing system consists of two identical TAT probes. When the aircraft is on the ground,the probes are aspirated by airfrom the bleed air manifold. Total Air Temperature Probe I Cabin Pressure Relief Air Valve: The TAT probe I cabin pressure relief air valve is a dual purpose valve which provides pressurized air to the TAT probes and the cabin pressure relief valve when the aircraft is on the ground. Operational Summary: Digital Air Data Computing: The primary function of the DADC system is to show altitude, airspeed and air temperature data on the primary flight display. Each MFP contains two sensors that measure P5 and Pr. Static pressureis the magnitude of the ambient air pressure around the aircraft while it is in flight. Ideal Ps does not contain pressure components from the forward movement of the aircraft. But, the air pressure that moves around the aircraft and the subsequent Ps port is slightly higher or lower because of the location of the port. This difference is referred to as the static source error. Total pressure is the air pressure measuredwith a tube that is open atthe front, closed at the rear and points directly into the airstream. It is the sum of static air pressure, impact pressure and total force for each unit area that is applied by air mass into the tube. Each MFP senses and calculates the absolute pressure that is applied to the P5 ports and Pr probes. The MFP calculates the following AD parameters: AOA Indicated Ps Corrected P5 Indicated Pr Corrected Pr Indicated impact pressure Corrected impact pressure Indicated dynamic pressure Corrected dynamic pressure Total air pressure SSEC Total source error correction The altitude parameters includes the following: Pressure altitude Barometric (Baro) correction Baro corrected altitude Altitude rate The airspeed parameters includes the following: Indicated airspeed Calibrated airspeed Indicated mach Calibrated mach True airspeed Maximum operating speed Maximum operating mach Overspeed warning The air temperature parameters include static air temperature. The maintenance parameters include fault messages. (1) Barometric Set Controls: The Flight Guidance Panel (FGP) has two BARO correction input knobs and two set to standard pushbuttons. The BARO knob transmits a count indicating relative knob movement to the Input I Output (1/0) modules via reflected binary code called grey code. The state of each set to standard pushbutton is also transmitted on the bus to the 1/0 modules. The FGP is connected to each modular avionics unit. Controls and Indications: (See Figure 2. Probe Heat Switches.) Static Source Error Correction (SSEC) Disable Switch: The SSEC disable switch is used to disable SSEC for DADC system testingon the ground. The SSEC disable switch provides a ground I open discrete into the MFP probes. Circuit Breakers (CBs): The flight environment data system is protected by the following CBs: Circuit Breaker Name CB Panel Location Power Source ADS 1CH 1A CPOP B-9 R ESS 28V DC Bus ADS 1CH 2A CPOP B-8 R ESS DC Bus ADS 2 CH 1B CPOP B-7 R MAIN 28V DC Bus ADS 2 CH 2B CPOP B-6 R MAIN 28V DC Bus ADS 3 CH 3A POP B-8 L Main DC Bus ADS 3 CH 4A POP B-9 L MAIN 28V DC ADS STBY CH 3B POP B-6 L ESS DC Bus ADS STBY CH 4B POP B-7 L ESS DC Bus AIR DATA PROBE 1 HTR PRI REER A-10 R ESS 28V DC Bus AIR DATA PROBE 1 HTR SEC REER B-10 R MAIN 28V DC Bus AIR DATA PROBE 2 HTR PRI REER A-11 R MAIN 28V DC Bus AIR DATA PROBE 2 HTR SEC REER B-11 R ESS 28V DC Bus AIR DATA PROBE 3 HTR PRI LEER A-2 L MAIN 28V DC Bus AIR DATA PROBE 3 HTR SEC LEER B-2 L ESS 28V DC Bus AIR DATA PROBE 4 HTR PRI LEER A-1 L ESS 28V DC Bus AIR DATA PROBE 4 HTR SEC LEER B-1 LMAIN 28V DC TAT PROBE 1 MCDU SSPC (#3415) LESS 28V DC TAT PROBE 2 MCDU SSPC (#3416) R ESS 28V DC Bus TAT VALVE MCDU SSPC (#3414) L MAIN 28V DC Bus c. Crew Alerting System (CAS) Messages: The following CAS messages are associated with the flight environment data system: CAS Message Posslble Causes ADS 1-2-3-Stby Fail (caution) Two or more of airdata systems failed. ADS 1-2-3-Stby Maint Reqd (caution) System fault in indicated air data system. OR AIR DATA PROBE heater switch(es) selected OFF while in flight. ADS 1-2-3-Stby SSEC lnvali (caution) Static Source Error Correction to indicated Air Data System is disabled or is unavailable and has reverted to the default table. Air Data Probe 1-2-3-4 Heat Fail (caution) Indicated air data probe heat failed. ADS 1-2-3 Stby Fai (advisory) Indicated Air Data System (ADS) has failed. ADS Miscompar (advisory) Priority FGC has detected an unflagged miscomparison between Air Data Systems. Limitations: Flight Manual Limitations: There are no limitations for the flight environment data system at the time of this writing. OEM Provided Data Navigation Revision 12 20225-19 2A-34-00: B of 36 Figure 1.Air Data System (ADS) Block Diagram TIL-002952 Figura 2. Probe Heat Switches SEE DETAILA AIR DATA PROBE ANTI ICE HTR SWITCHES Used to control electrical heating to the four Multi-Function and TAT probes. EVS WINDSHIELD HEAT WSHlDHT LF.RS RFILS BB CMIN WDOHT DETAIL A TIL-002948 2A-34-30: Standby Multifunction Controller (SMC) System 1. General: The purpose of the Standby Multifunction Controller (SMC) is as follows: Provide standby flight information OEM Provided Data Navigation Revision 12 20225-19 2A-34-00: 1a of 36 Provide primary display, Head-Up Display and Weather Radar control functionality Provide information and operation of utility functions The SMC serves multiple purposes withinthe Flight Deck. The primary role of the SMC isto provide standby flight information display. The secondary role of the SMC is to provide control of the primary DU- 1310 display units, the Head-up Display (HUD), and control of the RDR-4000 weather radar. The tertiary roleof the SMC is to provide various utilityfunctionality such as aircraft refueling capability, readouts from the hydraulic and oil quantity, and Semi-CPCS control.Pitot static information is provided by theAir Data System (Multi-function Probes). Attitude and heading information is provided by the aircraft Inertial Reference UnitsandtheAH-1000 Attitude I Heading Reference Source (AHRS).The AHRS unit interfaces with the magnetometer and the Air Data System. If the Display Valid discrete from any DU transitions from 'Valid' to 'Invalid', the SMC SFD shall be commanded to display, without crew action.Similarly, if an aircraft break-power transfer occurs during flight, the SMC SFD shall be commanded to display. Generally, the pilot's SMC controls the pilot's displays and the copilot's SMC controlsthe copilot's displays. However, there are some functions that can be selected from either SMC. Normal flight crew operation of the SMC requires the pilots to select the required display control menu via the SPU function keys. During normal operations, the SFDwill be crew selectable, but will not be required for display. During Ground Service operations, the SMC will acquire information from the FQMS and FQI for refueling and oil I hydraulic information. The SMC system consists of the following components: Two Switch Panel Units (SPUs) Two Display Panels (DPs) Two Attitude I Heading Reference System (AHRS) Two Magnetometer Units Description of Subsystems, Units and Components: Switch Panel Unit (SPU): (See Figure 3. SMC Switch Panel and Display Units.) The Switch Panel Units (2) provide the pilots with a means to select functions on the DP. There are 16 function keys in a 3x5 + 1format (three rows of five buttons plus one larger button for SFD display). The function keys contain internal annunciators to indicate to the pilots which item is displayed on the DP. When a function key is selected, the SMC will light the key's 'bar-type' annunciator and display the appropriate menu. During abnormal operations, specifically when the SMC is commanded to display the SFD, the annunciator for the Standby function key will be activated without pilot action. The function keys available on the SPU are as follows: PFD Display Control MAP Display Control 1/6 DU-1310 \Nindow Control 213 DU-1310 \Nindow Control Electronic Charts Control On-side Sensor Selection Flight Reference Functions Thrust Reference System Control Weather Radar Control Navigation System Control Test Control Checklist Functions HUD Control Utility Functions SFD Pop-up Display Panel (DP): (See Figure 3. SMC Switch Panel and Display Units.) The SMC DP is a self-contained Liquid Crystal Display (LCD), which is used to display the selected menus and SFD. The DP is comprised of a LCD, 10 line select keys, a range/set knob, and a brightness/dim knob which also acts as the standby menu activation mechanism.When the SMC is in display control mode, the system communicates with the MAU via ARINC 429 to acquire menu information, and to control MAU-based systems. The DP also communicates with non-MAU based systems directly via ARINC 429. The range/set knob is a dual-purpose knob which controls the on-side Map display range and provides value-set functionality. The range/set knob also provides value-set functionality in the same manner as previous DC-884 Display Controllers. The brightness/dim knobinternally controls the brightness of the LCD display. The standby menu activation knob will communicate internal to the SMC to display the SFD menu. Power for the SMC DP is supplied by the Emergency Battery DC Power bus or the Ground Service DC Power bus. Power is supplied to each DP through separate circuit breakers. Attitude I Heading Reference System (AHRS): (See Figure 4. Typical Heading and Attitude Displays on SMC.) The AH-1000 AHRS unit is a stand alone attitude and heading reference source. The AH-1000 interfaces with the magnetometer, which acquires heading and attitude information. The output from the AH-1000 is a two-output, single-source ARINC 429 signal to each SMC DP. Power for the AH-1000 is supplied by the Emergency Battery DC Power bus. Controls and Indications: Circuit Breakers (CBs): The following circuit breakers power the Standby Flight Instruments: Circuit Breaker Name CB Panel Location Power Source AHRS 1 POP A-3 L EMER DC Bus AHRS 2 CPOP A-3 R EMER DC Bus MAGNETOMETER 1 POP A-2 L EMER DC Bus MAGNETOMETER 2 CPOP A-2 R EMER DC Bus SMC 1 POP C-3 L EMER DC Bus SMC 1(GND) MCDU SSPC (#3122) R MAIN DC Bus SMC 2 CPOP C-3 R EMER DC Bus SMC 2 (GND) MCDU SSPC (#3123) R MAIN DC Bus Crew Alerting System (CAS) Messages: The following CAS messages are associated with the Standby Flight Instruments: CAS MesHge Possible Causes @Wlfjtft l (advisory) Indicated SMC has failed. Limitations: There are no SMC limitations at the time of this writing. Figure 3. SMC Switch Panel and Display Units BRT (Brightneu) Knob Rotating the BRT knob increases (CW) or decreases (CCW) the display brightness Whan the SMC is in standby mode, depressing the M on the BRT knob activates the Standby menu (as depicted in the above display) TIL-004116 Figura 4. Typical Heading and Attitude Displays on SMC SEE DETAIL A ATTITUDE (AHARS) HEADING (MAGNETOMETERS) DETAIL A TIL-002949 2A-34-40: Head Up Display (HUD) II System 1. General Description: The Gulfstream Head Up Display II (HUD II) is the Rockwell Collins Model 6250 Head-up Guidance System (HGS) which is an electronic and optical Head Up Display (HUD) system that generates and displays information in the pilot's forward field of view. The displayed infonnation is derived from the aircraft instruments and sensor data. The HUD II provides information during all phases of flight. The HUD IIsupplies attitude, speed, altitude, flight path, Flight Director guidance,Traffic Alert and CollisionAvoidance System (TCAS) guidance, visual approach, unusual attitude, and other situational infonnation to the pilot in symbolic format. The HUD II installation contains a Pilot Display Unit (PDU) that consists of Combiner and an Overhead Unit (OHU), driven by a HUD II Computer (HC). Airplane specific HUD installation data are stored in an off-the-shelf HUD Personality Module (HPM). HUD II also has the ability to display data from the Enhanced Vision System (EVS),a thennal image of the realwortd scene invideo fonnat to enhance the pilot's situational awareness at night andin certainlow visibility environmental conditions. This image is generated by an infrared sensor (camera),which is installedin the nose of the airplane between the radome and windshields. The infrared image is scaled, aligned and adjusted (as required)to overlay the real world scene as viewed through the Combiner. EVS image adjustments are also stored in the HPM. The HUD II encompasses the functions of four components: The Combiner Assembly The Overhead Unit The HUD II Computer The HUD Personality Module (HPM) Description of Subsystems, Units and Components: Combiner Assembly The Combiner (see Figure 5. HUD Combiner) is an optical element that combines displayed flight symbology with the pilot's view through the airplane's windscreen. This optical design permits various symbols, such as the artificial horizon, to overlay corresponding features of the outside world. The Combiner,which does not require any alignment or adjustments after installation, is installed on the forward windscreen frame in the flight compartment so that the HUD II display is viewed from the pilot's nonnal seated position. The Combiner display is focused at optical infinity, allowing the pilot to see the infonnation without a shift in eye focus. The control panel on the Combiner is on the inboard side, which allows the pilot to use the inboard hand to control changes. The pilot has the ability to adjust display brightness, select Auto or Manual Brightness Mode, and adjust video contrast and brightness on the Combiner control panel.The manual adjustment mode allows the pilot to set a constant luminance. The automatic adjustment mode allows the pilot to set a constant contrast ratio. Actual luminance of the Combiner display varies with the sensed luminance of the ambient light. Ambient light sensor signal provides feedback to the HC to automatically adjust brightness levels based on ambient lighting conditions if the brightness mode is set to Auto. Video brightness and contrast are controlled independently from each other by means of two separate potentiometers on the Combiner. Brightness and contrast controls have positive stops at the minimum and maximum positions and have 270° ±5° of travel.The brightness control knob can be used to adjust the display brightness from zero to full intensity. The Combiner sends signals to the OHU for display brightness/contrast control settings, stow switch setting, ambient brightness sensor reading, Combiner alignment position detedor reading, and Combiner strapping discretes. The Combiner has a stow mechanism that allows it to be stowed when not in use. The pilot simply removes the Combiner from the forward field of view by moving it to the stowed position. The Combiner receives electrical power from the OHU. The Combiner does not require forced air-cooling and weighs approximately 5.0 lb. (2.3 kg). Overhead Unit The Overhead Unit (see Figure 6. HUD Overhead Unit and Combiner Installation) projects the flight symbology, generated by the HC, onto the Combiner. VVithin the OHU is an Adive Matrix Liquid Crystal Display (AMLCD), a relay lens assembly, backlight and supporting eledronics. The OHU employs a folded optical path design using a prism. This allows the assembly to be asfar forward and as high as possible inthe flight deck to maximize head dearance. The OHU receives the pixelated image data, consisting of merged EVS video and flight symbology, from the HC. The display image light passes through the OHU optics and is relayed to the Combiner, where the light rays are aligned and finally viewed by the pilot. The OHU receives signals from the HCfor displaying the symbology, EVS video images, display brightness, HUD failure, and the LED Enable BIT. The OHU receive signals from the Combiner for display brightness control settings, contrast setting, stow switch setting, ambient brightness sensor reading, Combiner alignment position detedor reading, and Combiner strapping discretes. The OHU transmits signals to the HC for display feedback, BIT data, and Combiner strapping data. The OHU is powered by the HC. The electrical components of the OHU are cooled by forced air provided by integral fan and therefore does not require external forced air cooling. The weight of the OHU is approximately 27.0 lb (12.2 kg). HUD II Computer The HUD II Computer (HC) provides a pixelated image to the OHU for display via the Combiner. This image is derived from internal calculations, which determine the position and form of the flight symbology to be presented, dependent on the airplane's equipment and sensor data received by the HC. Additionally, the HUD II Computer evaluates the performance of the HUD II by monitoring the state of the HUD II and airplane inputs. The HCis powered by the airaaft 28V DC bus and weighs approximately 14.0 lb. (6.4 kg). HUD Personality Module (HPM) The HPM is used to store airplane specific HUD installation data such as electronic boresight, EVS offsets and airplane specific data, outside of the HUD II computer. This allows for removal and replacement of the HC or PDU without having to manually reenter this data. The HC p01Ners the HPM. All datatransmissions are between the HC and HPM. The HPM weighs approximately 0.1 lbs (0.05 Kg). Video Processing The HGS can display real time video images from the Enhanced Vision System (EVS). The video images are overtayed with the HGS symbology before they are displayed on the Combiner. Display of the video image is controlled by the pilot via a selection on the Standby Multifunction Controller (SMC) and an associated button on the control yoke. Operational Summary: The HUD II provides attitude, speed, altitude, flight path, Flight Director guidance, Traffic Alert and Collision Avoidance System (TCAS) guidance, visual approach, unusual attitude and other situational information to the pilot in symbolic format. Basic flight information is available during all phases of flight. The HUD also displays a video image from EVS. The HUD II display representation of airaaft sensor and equipment data is designed for immediate recognition by the pilot while not causing any confusion due to ambiguity with like data presented on other cockpit displays. The pilotis made aware ofthe loss of any required HUDdisplayed information, due to unavailable aircraft sensor or equipment data, through the obvious blanking of display elements or the display of special status messages. The SMCs provide control and display lines for entry of HUD reference data and selections of EVS and Synthetic Vision System (SVS) settings. When flying an approach usingthe HUD, the approach descent angle andthe runway elevation can be entered to provide improved awareness to the pilot. The descent anglewill provide a reference line for the FPV for alIapproaches, and the runway elevation will ensure the proper perspective is applied to the synthetic runway display for ILS or LPV approaches. If conducting an ILS or LPV approach using QFE settings, the runway elevation shouldbe set to zero. The HUD II display brightness can be adjusted from low to full intensity through the brightness control knob located on the Combiner. A manual adjustment modeis availablethat allows a constant luminance to beset bythe OEM Provided Data Navigation Revision 12 20225-19 2A-34-00: 1B of 36 pilot. An automatic adjustment mode is also available that allows a constant contrast ratio to be set by the pilot (the actual luminance of the Combiner display varies with the sensed luminance of the ambient light). The Combiner is removed from the pli ot's forward field of view by moving it to the stowed position. VVhen selected, the thermal video image generated by the EVS sensor will be displayed on the Combiner, merged with the HUD II symbology. There are three gain settings for the thermal image: automatic gain, high gain and low gain. The default selection is automatic gain, and this setting automatically adjusts the contrast of the image based on ambient conditions. The high gain setting displays the image at a high contrast ratio, while the low gain setting displays the image at a low contrast ratio.There are separate brightness and contrast controls on the Combiner to adjust the EVS image. The HUD records the detected faults (in airplane's equipment and sensors, and the HUD LRUs) and sends this data to the central maintenance computer (CMC). The HUD provides electronic boresight adjustment of up to 10 milliradians laterally and 10 milliradians vertically to compensate for installation offsets.These adjustment values (called Boresight offsets) are stored in the HPM. The HUD also provides the ability to adjust the EVS image to compensate for different camera tolerances and installation offsets. These adjustment values (known as EVS offsets) are also stored in the HPM. Note Video from the HUD computer is routed to Video Module 2 only. This allows a combined HUDI EVS image to be displayed on DU 3 and DU 4. EVS Video is sent to Video Module 1and Video Module 2, providing EVS-only imagery on all DUs. Controls and Indications: Circuit Breakers: The following circuit breakers protect the Head Up Display II (HUD II): Circuit Breaker Name CB Panel Location Power Source HUD FANNIDEOAMP POP B-1 Left Main DC HUD PWR 1 POP B-2 Left Main DC HUD PWR 2 CPOP B-2 Right Main DC Crew Alerting System (CAS) Messages: The following CAS messages are associated with the HUD II: CAS Message Malfunctlon/Fault HUD Comp Fan Fail (advisory) Indicates the cooling I ventilation fan for the HUD computer has failed. HUD OHU Fan Fai (advisory) Indicates the cooling I ventilation fans in the Overhead Unit have failed. Built-In Test Equipment (BITE): The HGS contains extensive BIT capabilities to detect and localize faults. The BIT capabilities are: Power-up Self-Test - A series of tests performed automatically each time the system is powered up. Continuous (background) BITE - Following completion of Power-up Self-Test, a set of BITE checks are continuously performed in the background. These tests continuously monitors the main (criticaO processing path of the HC, and the display signal routing and feedback paths to and from the Overhead Unit. These tests do not affect normal (no fault) operation of the HGS. No single power supply failure can disable the continuous BITE. Operator-Initiated BITE - Operator-Initiated BITE consists of a set of tests that are manually initiated when the aircraft is on the ground, and are used to perform comprehensive tests on the system. These tests can be initiated from the Central Maintenance Computer (CMC). HGS System Monitoring Display Path Monitoring- The HGS monitors the display path to prevent the display of misleading data. Display path monitoring is done by enabling a Critical Symbol Monitor (CSM) and by over1apping the BITin the OHU and HC. The CSM tests the integrity of the HGS display path by monitoring the displayed position of selected critical symbols. Maintenance Monitor - The HGS continuously gathers data for the detection and identification of HGSinternal faults and faults of systems thatinterface with the HGS.The status andfailurecodes are sent to the aircraft CMC. Limitations: Category IHUD Operations: Category IHUD operations are approved. Non-Directional Beacon Approaches: The HUD does not provide a Non-Directional Beacon (NDB) approach capability. NDB approaches may be set up and flown through the FMS, using the HUD for guidance. Figure 5.HUD Combiner BORESIGHT HORIZON LINE AIRPORT SYMBOL MARKER BEACON CONFORMAL LATERAL DEVIATION SCALE CONFORMAL VERTICAL SPEED READOUT DIGITAL RADIO ALTITUDE LATERAL DEVIATION CENTER LINE DECISION HEIGHT ANNUNCIATOR TIL-002951 Figure 6. HUD Overhead Unit and Combiner Installation SEE DETAILA SEE DETAIL B COMBINER ASSEMBLY OVERHEAD UNIT DETAIL A CONTR MAN 0 ® MIN AUTO VIDEO @ HUD BRT BRT QJ MIN MIN '- ,,, DETAIL B TIL-002950 2A-34-50: Enhanced Vision System (EVS) II GeneralDescription: The EVS II system consists of a Forward Looking Infrared (FUR) Camera integrated with the Visual Guidance System (VGS)to providethe airplanewith the capability to perform low visibility approaches to a 100 ft. decision point by enabling the pilot to see and identify the visual references identified in TiUe 14 CFR Part 91.176. This is accomplished by projecting the FUR image of the runway environment onto the Head Up Display (HUD) in a high-resolution raster format, such that the image is conformal with the outside visual scene. EVS II is functionally equivalent to the previously certified EVS. EVS II includes fewer system components as well as smaller and lighter components. EVS II is composed of the following components: Kollsman Forward Looking Infrared (FUR) Assembly Processor Assembly EVS 11 Window Assembly The EVS has no control panel of its own.This function is performed by the Standby Multifunction Controller (SMC) (see Figure 7. EVS Menuon the SMC). Description of Subsystems, Units and Components: Field of View (FOV): The EVS II infrared camera field of view (FOV) is a nominal 30.0° (+0.5°/-1.0°) horizontally and 22.5° (+1.0°/-0.5°) vertically. Imagery captured by the camera is routed through a processor which converts the infrared picture to a raster format and then communicates the formatted data to the HUD overhead unit. The HUD overhead unit synchronizes the data with the navigational and situational symbology presented on the HUD combiner to display a congruent image with an extended visual range. Forward Looking Infrared (FUR) Assembly: The FUR Assembly, i.e., EVS II camera, (see Figure 8. FUR Assembly Installation) is mounted above the radome (forward of the windshields) in order to provide the EVS II camera a clear, unobstructed forward field of view parallel to the airplane's horizontal reference line. The unit produces both a composite video signal and digital video data. The unit containsthe Sensor Assembly, which is an Infrared Camera, Video Card and Auxiliary Card.The infrared video image produced bythe FURAssembly is sent to the Processor Assembly line replacement unit (LRU) and then ultimately projected onto the HUD. Processor Assembly: The Processor Assembly is the control and monitoring element of the EVS IIsystem. It provides a functional interface between the HUD andthe FUR Assembly. Video is output in both analog and digital format, which are then displayed on the pilot's HUD in a 1:1 overtay to the scenery viewed through the windshield. The Processor Assembly sends and receives signals with the airplane systems via the ARINC-429 communication bus. These communications include control signals, status, and operational data, commands to tum the FUR Sensor Assembly on and off, initialization of Built-In-Test (BIT), changes to FLIR Sensor gain, and to initiate the FUR Sensor Non-Uniformity Correction (NUC) process. Status messages indude: Results of FLIR BIT and NUC processes FLIR Assembly status Temperature measurements with in the FUR Assembly Video Validity Processor and FLIR operating times EVS II V\findow Assembly: The EVS IIwindow assembly consists of a sapphire windowwith a heater grid deposited on the inner surface of the window and anti-reflection coatings on both inner and outer surfaces. It allows the FLIR to view the scene while affording protection from the elements. WindCJIN Heater Control: (See Figure 9. EVS Windshield Heat Switch.) The V\findow Heater is controlled by the Processor Assembly. It applies 28V DC power to the heater grid on the window as needed to remove condensation or remove ice buildup on the windCJIN. Should anything prevent the heater grid from heating,the Processor Assembly sends an EVS WINDOW HEAT FAULT to the CMC. When the airplane's anti-ice system is active, heat is applied to the window in two different methods: With the gear UP, the heat cydes one minute on, seven with thermostat minutes off. When landing gear is down, the heat is ON continuous with thermostat. Continuous heat with thermostat: This operation starts when one of the heater "ON" discretes is selected and the temperature of the EVS window is less than 12°C (54°F), allowing power to the heating grid. When the temperature reaches 14°C (57°F), power to the heating grid is removed. Thermostat protection triggers if the window temperature exceeds 70°C (158°F). An (advisory) CAS message is generated, and a fail condition is generated in the window heat status page of the CMC EVS page. The following table summarizes EVS IIwindow heat logic: Left Cowl Anti-Ice Right Cowl Anti-Ice Landing Gear Position Window Heat CMD ON EVS Window Heat Appllcatlon No ice No ice x VVindow Heat Off Heat OFF No ice Ice Up VVindow Heat Off Heat ON - Cydes (7 minutes OFF, 1minute ON) Left Cowl Anti-Ice RightCowl Anti-Ice Landing Gear Position Window Heat CMD ON EVS Window Heat Application Ice No ice Up Window Heat Off Heat ON - Cycles (7 minutes OFF, 1 minute ON) Ice Ice Up Window Heat Off Heat ON·Cycles (7 minutes OFF, 1 minute ON) No ice Ice Down Window Heat Off Heat ON - Continuous Ice No ice Down Window Heat Off Heat ON • Continuous Ice Ice Down Window Heat Off Heat ON - Continuous x x x Window Heat on Heat ON·Defrost (2 minutes ON, then OFF) limited by the window thermostat 14°C {57°F) Anti-ice discrete inputs are Open = Ice and 28V = No Ice. Also, the heat can be applied manually by the pilot for two minutes using the switch on the overhead panel Qimited by the window thennostat 14°C (57°F). f. Environmental Operating Charaderistics The table below provides the environmental charaderistics of the EVS II components. Parameter FLIR Processor IR Wlndow Operating Low Temperature .55•c 15 c .55•c Operating High Temperature +70°C +10°c +10°c Short-Time Operating High Temperature +70°C +1o·c +1o c Ground Survival Low Temperature .55•c .55•c .55•c Ground Survival High Temperature +85°C +85°C +B5°C Altitude 55,000 ft. 15,000 ft. 55,000 ft. Decompression N/A 8000 TO 55,000 ft. MSL N/A Overpressure N/A -15,000 ft. N/A Power Source 28 Vdc 28 Vdc 28 Vdc Weight (MAX) 12 lbs (5.44 kg) 10 lbs (4.53 kg) 2.5lbs (1.13 kg) Humidity 85% 85% 100% 3. Operational Summary: Power On Operation When power is applied to the EVS II system and the airplane is on the ground, a Power-up BIT (PBIT) is performed,taking approximately fifteen seconds. When power is appliedin the air, the EVS IIskips the Power-up BIT and completes initialization within five seconds, given that PBIT is performed prior to takeoff in order to be powering up in air. Then an automatic Built-In-Test (BIT) sequence is initiated; all system functions are checked. The BIT takes approximately one minute to go through its cyde. If a fault is discovered by the BIT during power-on, an EVS FAILURE, EVS MAINTENANCE REQUIRED or EVSWINDOW FAILURE message appears on the CAS display. otherwise, when BIT is completed, FUR is annunciated on the HUD, indicating the FUR is ready for operation. When initially powered, the EVS II FUR may take up to fifteen minutes to cool its detector to operating temperature and activate for an initial power-on. If the EVS has been previously activated, this activation period may be shorter depending on the deviation from nominal operating conditions that has occurred since power-off. When the aircraft is parked overnight in temperatures below minus forty degree centigrade (-40°C), the EVS II FUR may take up to 35 minutes to heat the electronics to operating temperature and to cool the detector to operating temperature. The EVS II Processor will output video from the FUR duringthe cool-down process. However, this video is only displayed on the head-down display during this time. Once the FUR is cooled to temperature, the EVS IIwill perform a NUC and then enable video to be displayed on the HUD.The NUC completes its cycle within one minute. Normal Mode Operation The EVS acquires its imagery with the FUR Assembly. The video is sent to the processor and forwarded from there to the VGS. The image activity is projected onto the HUD combiner glass in a combined stroke I raster format. The imagery is combined with the necessary flight guidance and navigation data normally displayed on the HUD combiner glass when the EVS is not in use. The combiner glass is deployed or stowed using a release lever on the combiner sidearm. A brightness control knob and a switch to select either Manual or Normal mode is located on the combiner assembly. When the mode switch is in Manual mode, the display brightness is adjustable from zero to full intensity using the brightness control knob. When the switch is in the Normal mode, the display automatically adjusts with ambient light changes in the cockpit detected by the ambient light sensor.The ambient light sensor is mounted on thefront of the combiner facing the windshield. Limited adjustments to brightness are allowed in the Normal mode using the brightness control knob. An EVSNGS clear button on the yoke functions as an EVS video toggle button. If pressed while EVS video is displayed on the HUD, EVS video will be immediately removed from the HUD. This button also affects VGS guidance data displayed on the HUD. SubsequenUy pressing this button will toggle EVS and VGS video feeds back ON. Ifthe EVS is OFF (has not been through the power-on sequence) or the EVS has been deactivated since the EVSNGS dear button was pressed, this button will not affect (activate) the EVS. Maintenance Mode Operation Maintenance mode is available only when the Weight On Wheels CNQW) discrete value is true and the Maintenance Test Switch is ON. Maintenance mode allows for adjustment and operational assessment of the EVS components. The Central Maintenance Computer (CMC) is used for maintenance information. While in this maintenance mode, registration of the EVS image is done through this CMC interface. Built-In-Test (Bit) Mode Operation When in BIT mode, EVS II cannot be used for other activities. The entire system is occupied with test functions. BIT mode is initiated at power-on and at the user's discretion. Results are displayed on the maintenance pages. The exception iswhen a maintenance-required warning or an EVS failure is reported. These warnings are displayed on the Crew Alerting System (CAS). The reasons for these warnings can usually be ascertained from the CMC pages where system status information reveals abnormalities. Continuous On-line BIT is not part of the BIT mode. It is a continuous and non-intrusive operation that runs in the background.A limited number of checks are made in this way. The power-on BIT and manually initiated BIT are more extensive, and as a result, when they are in progress they take over the continuous on-line BIT. The initiated BIT is started through the CMC. Alignment And Registration Boresighting of the EVS Camera Mount is performed with reference to the installed and aligned VGS. The EVS is successfully registered when the EVS video depicting the real world accurately overlays the actual real wor1d view through the cockpit window. Controls and Indications: Circuit Breakers: ...-C-a.t- -n .- . IF AN EVS II/ HUD SYSTEM FAILURE RESULTS IN A CIRCUIT BREAKER POPPING IN FLIGHT, DO NOT RESET THE CIRCUIT BREAKER. The following circuit breakers power and protect the EVS II system: OEM Provided Data Navigation Revision 12 2022-05-19 2A-34-00: 'll of 36 Circuit Breaker Circuit Breaker Panel Location Power Source EVS CAM MCDU SSPC (#3404) L MAIN DC Bus EVS HTR MCDU SSPC (#3405) L MAIN DC Bus EVS PROC MCDU SSPC (#3406) L MAIN DC Bus Crew Alerting System {CAS) Messages: The following CAS messages are associated with the EVS II: CAS Message Malfunction/Fault 1441ttJI {caution) Indicates serious failure within the EVS system. The EVS system is shut down and the pilot is not allowed to use it. Any imagery displayed on the HUD is immediately removed. EVS Window Heat Fail (caution) Indicates a failure of the EVS window system. The EVS remains operational at the discretion of the pilot (although this should be avoided under icing conditions). 1·•JI8·• {Mad&v)is·Rory•) Indicates maintenance requirement. The message occurs only if there is a non-critical maintenance issue with EVS that should be addressed by maintenance personnel. When a maintenance message is issued, the continued use of EVS is left to the discretion of the pi ot. Other Annunciations: Annunciations of hazardous conditions are displayed on the HUD combiner when sensed by the Enhanced Ground Proximity warning System (EGPWS). See ACM Chapter 28 for a tabulation of EGPWS hazard text messages. Flight Manual Limitations: Pilofs Manuals: Usingthe EVS requires the 2A-34-40: Head Up Display (HUD) II System and 2A-34-50: Enhanced Vision System (EVS) II System Descriptions to be immediately available to the flight crew. Presence of Visual Cues: At 100 feet HAT, visual cues must be seen without the aid of EVS to continue descent to landing. Qualifications For Use: EVS may be used only by qualified pilots who have been trained in accordance with requirements listed in the FAA GVI Flight Standardization Board (FSB) Report. Vertical Guidance Requirements For IMC EVS Aproaches: Flight Director or autopilot with vertical guidance, either ILS or FMS vertical path, is pennitted for all IMC EVS approaches. EVS Requirements: EVS, as installed, meets the requirements of EFVS (Enhanced Flight Vision System) as defined in 14 CFR Part 91.176. Figure 7. EVS Menu on the SMC EVS SELECTIONS ON HUD MENU NUC: (Non-Uniformity Correction): Selects NUC ON and OFF. Default is OFF FLIR: Selects FLIR camera ON and OFF, also selects EVS OFF (if currently selected ON). NOTE: May take up to 30 minutes to appear when FUR is selected ON. EVS: Selects EVS ON and OFF. Selecting EVS ON also selects FLIR camera ON (if currently selected OFF). AUTO I HL: Toggles FUR camera through AUTO, high (H) and low (L) gain settings. Default is AUTO. TIL-002517 Figure 8. FUR Assembly Installation FLIR ASSEMBLY Provides a composite video signal and digital video data. Contains the sensor assembly (infrared camera), video and auxiliary cards. TIL-002515 Figure 9.EVS Windshield Heat Switch SEE DETAILA ANTI ICE HTR WINDSHIELD HEAT EVS - AIR DATA - O WSHLD HT PROBE 1 PROBE 2 LF I RS RF I LS CABIN WOO HT PROBE 3 PROBE 4 0 0 DETAIL A TIL-002516 2A-34-60: PlaneView II- Block 2 Installation General Description: This section describes the operationally significant changes that were induded in the PlaneView II Avionics Software version "Block 2" upgrade, installed with Aircraft Service Change (ASC) 902. All G650 I G650ER aircraft in service have Block 2 installed. System Information: 1. CAS Messages: To provide consistency of terminology and ease interpretation, two GAS messages that used the term "rudder steering" were changed to more appropriately refer to nose wheel steering through the useof rudder pedals as "pedal steering". The Rudder Steering Fail (caution) CAS message was changed to (caution), and the Rudder Steering Off (advisory) CAS message was changed to {advisory). The AP Control Switch Stuck (advisory) CAS message was removed. It was redundant with the remaining {advisory) CAS message. A one-second delay was inserted into the logic for autobrakes mode CAS message display. VVhen an autobrakes mode selection is changed (e.g., pilot selects MED and then changes to LOVV), there is a small amount of time required for the Brake Control Unit to process the change. This often resulted in a momentary display of both autobrakes mode GAS messages before only the final selected mode CAS message remained.This one-second delay for displaying the selected autobrakes mode CAS message significantly reduced the occurrence of two autobrakes mode CAS messages momentarily displaying simultaneously. The (caution) CAS message logic was changed to include a tire overspeed event. Note that this GAS message is not retained following removal of all electrical power from the airplane; therefore, it is imperative to record the information in the maintenance log when the event occurs. 2. Autopilot Disengage Indication on PFD: Regulatory requirements mandate that a "warning" (i.e., red visual indication) must be provided following autopilot disengagement, regardless of cause. Prior to installation of ASC 902, a pilot-initiated manual autopilot disengagement resulted in an amber indication onthe PFD, and an automatic AFCS autopilot disengagement (caused by a fault or failure) resulted in a red indication. With ASC 902 installed, the autopilot disengage indication on the PFD is red, regardless of the cause. Datalink Weather Coverage Expansion: The graphical NEXRAD weather uplink capability was improved to provide wor1d-wide coverage. Note that while a weather request can be submitted, receipt of graphical weather from remote areas of the world may be limited. DME Hold Function: The "Block Point 1" avionics upgrade errantly inhibited DME Hold capability with FMS selected as the navigation source. With ASC 902 installed, DME Hold can be selected regardless of the navigation source in use. Flight Controls 213 Synoptic Paga: The Flight Controls 213 synoptic pageformat (detailed or decluttered) isstored in non-volatile memory, and the last displayed format is recalled for use the next time the Flight Controls 213 synoptic is selected for display. Additionally, when the Flight Controls 213 Synoptic page is selectedfor display via the CCD, the cursor position defaults to the "FCC - Actuators" selection box to provide a rapid means of changing display formats. Flight Director Flight Level Change (FLCH) Mode Transition: With the autopilot engaged, the Flight Director FLCH mode automatically transitions to PIT modewhenthe autopilot is disengaged. The objective of this logic change was to provide speed control with the autothrottles, if engaged, inthe event the pilot does not satisfy the Flight Director vertical guidance after autopilot disengagement. If the pilot desires to continue the descent using FLCH mode after autopilot disengagement, the FLCH mode is available for re-selection. Inthe event FLCH is the Flight Director vertical mode and no lateral mode is selected (i.e.,ROL mode is displayed as the lateralmode on the PFD Flight Guidance Annunciation display), if the autopilot is disengaged the Flight Director cue will be removed and there will be no lateralor vertical modes indicated in the PFD Flight Guidance Annunciation display because PIT and ROL are considered standby modes. In an emergency descent with the autopilot engaged in AFCS Emergency Descent Mode (EDM), the lateral mode is ROL even though EDM is displayed in the Flight Guidance Annunciation; therefore,if the autopilot is disengaged duringthe descent prior to ASEL or ALT capture, the Flight Director cuewill be removed and there will be no lateral or vertical modes indicated in the PFD Flight Guidance Annunciation display. FMS Takeoff and Landing Data (TOLD): A revised TOLD database induded autobrakes performance data when landingwith an autobrakes mode selected.There is no performance credit for usingthe autobrakes RTO mode selection for takeoff; therefore, no autobrakes mode selection is available in the FMS TOLD Takeoff Initialization pages. VVhen landing with an abnormal flaps setting and/or on a contaminated runway surface, the only autobrakes mode selection available in TOLD is HIGH; however,landing data for authorized configurations are located in the appropriate G650 I G650ER Airplane Flight Manual Supplement for Autobrakes installation (G650-2013-10 and G650ER-2014-03). Additionally, ASC 902 induded multiple minor corrections made to FMS TOLD implementation. Forward andAft Emergency Batteries Voltages on DC Power 2/3 Synoptic Page: The Forward and Aft Emergency Batteries voltages are displayed in separate indication boxes on the DC Power 213 synoptic page, providingthe crew the capability to readily acquire the Forward and Aft Emergency Batteries charge state. Hydraulic Aux Pump Auto-Activation with Brake Application: Vvtth ASC 902 and ASC 051 (SPDS Build 10software upgrade) installed, if the hydraulic aux pump automatic function is armed (AUX PUMP OFF/ARM switch depressed), depressing a brake pedal with the aircraft on the ground and with low inboard brake accumulator pressurewill cause the hydraulic aux pumpto automatically activate, pressurizethe inboard brake accumulator, and then deactivate. 1-NAV MAP VSD Reverts to Track Mode when Autopilot Disengaged: Vvtth FMS lateral mode captured and the autopilot engaged, the 1-NAV MAP VSD will be in Flight Plan mode; however, when the autopilot is disengaged, the VSD will revert to Track mode. Conversely, with FMS lateral mode captured, the VSD will transition from Track mode to Flight Plan mode when the autopilot is engaged. Incremental Terminal Charts Database Loading: V\/hen loading a new Terminal Charts database, a comparison with the currently loaded database is automatically performed and only those terminal procedures charts that have been modified will be loaded I deleted. Lightning Sensor System CAS message logic: The "Block Point 1" avionics upgrade induded an implementation error that resulted in the LX Fail(caution) GAS message being displayed when the Lightning Sensor System (LSS) was operating normally. ASC 902 corrected this error, allowing the LSS to be operated by the flight crew. Flight Data Recorder Activation Logic: Prior to installation of ASC 902, the Flight Data Recorder (FDR) activated (i.e., initiated recording) only after the right-hand engine completed a successful start. Because there can be important data that could be utilized to investigate an abnormal engine start event, ASC 902 revised the FDR activation logic to beginrecordingwhen either the MASTER START switch or MASTER CRANK switch are selected to ON.