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Typical electronic

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ACARS,TCAS,IRS,FMS.

ACARS,TCAS,IRS,FMS.

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  • 1. Typical electronic/ Digital Aircraft System
    1
  • 2. Typical electronic/ Digital Aircraft System
    • ACARS - ARINC Communication and Addressing and Reporting System
    • 3. ECAM - Electronic Centralized Aircraft Monitoring
    • 4. EFIS - Electronic Flight Instrument System
    • 5. EICAS - Engine Indication and Crew Alerting System
    • 6. FBW - Fly by Wire
    • 7. FMS - Flight Management System
    • 8. GPS - Global Positioning System
    • 9. IRS - Inertial Reference System
    • 10. TCAS - Traffic Alert Collision Avoidance System
    2
  • 11. ACARS - ARINC Communication and Addressing and Reporting System
    The aircraft communications addressing and reporting system (ACARS) is a data link communication system. It lets you transmit messages and reports between an airplane and an airline ground base.
    3
  • 12. ACARS - ARINC Communication and Addressing and Reporting System
    These are typical ACARS reports:
    • Crew identification
    • 13. Out, off, on, in (OOOI) times
    • 14. Engine performance
    • 15. Flight status
    • 16. Maintenance items.
    4
  • 17. ACARS - ARINC Communication and Addressing and Reporting System
    These are the components of the ACARS:
    • Interactive display unit (IDU)
    • 18. Control display unit (CDU)
    • 19. ACARS program switch modules
    • 20. Management unit (MU).
    • 21. One can use the control display unit (CDU) or the IDU to control the operation of the ACARS and the CDU to show ACARS messages. The ACARS program switch modules contain dual inline package (DIP) switches. These switches identify the airplane.
    • 22. The ACARS MU receives the ground-to-air digital messages (uplink). It controls the transmission of the air-to-ground digital messages (downlink).
    5
  • 23. ACARS - ARINC Communication and Addressing and Reporting System
    ACARS connects to these other systems components:
    • VHF transceiver - to transmit to and receive data from the ground.
    • 24. Printer - to print ACARS reports and messages.
    • 25. Remote electronics unit - to distribute the chime annunciation and
    light annunciation signals.
    • Audio control panel - to signal the flight crew of incoming ACARS messages requiring flight crew attention.
    • 26. SELCAL control panel - to signal the flight crew of incoming ACARS message that require flight crew attention.
    • 27. Proximity switch electronics unit - to send discrete signals for out, off, on and in (OOOI) events.
    6
  • 28. ACARS - ARINC Communication and Addressing and Reporting System
    ACARS also connects to these systems to upload information from the airline operations or download information to the airline operations:
    • Flight management computers
    • 29. Flight data acquisition unit
    • 30. Data loader control panel
    7
  • 31. ACARS - ARINC Communication and Addressing and Reporting System
    8
    Figure : ACARS Management Unit and Interface
  • 32. ACARS - ARINC Communication and Addressing and Reporting System
    Management Unit
    The ACARS management unit (MU) gets uplink data from the VHF transceiver. It controls downlink data transmission from the VHF transceiver.
    The MU processes only uplink messages that have the airplane registration code. This same code is sent on all downlink messages to identify the airplane.
    ACARS Programming
    The ACARS program switch modules contain dual inline package (DIP) switches. These set the airplane identification and registration codes.
    9
  • 33. ACARS - ARINC Communication and Addressing and Reporting System
    Operation
    The flight management computer system control display unit (CDU) or the interactive display unit (IDU) gives one an interface with the ACARS system. They lets one enter, send, and review downlink/uplink data. The CDU shows ACARS messages in the scratch pad, the IDU shows ACARS messages on the display.
    IDU Features
    The IDU has a MENU key that selects the menu. From the menu page, you can select ACARS or the flight data acquisition unit aircraft condition monitoring system (ACMS).
    Software Loading
    10
  • 34. EFIS – Electronic Flight Instrument System
    General Description:
    A complete EFIS installation in an aircraft is made up of left (Captain) and right (co-pilot) systems, each system in turn being composed of two display units (an attitude director indicator (ADI) and a horizontal situation indicator (HSI), a control panel, a symbol generator and a remote light sensor. A third (centre) symbol generator is also incorporated so that drive signals from this generator may be switched to either the left or right display units in the event of failure of their corresponding generators.
    11
  • 35. EFIS – Electronic Flight Instrument System
    12
  • 36. EFIS – Electronic Flight Instrument System
    13
    Figure : Simplified block diagram of display unit
  • 37. EFIS – Electronic Flight Instrument System
    Display unit
    • The display units consist of the following chassis-mounted elements: a low-voltage power supply, a high-voltage power supply, four circuit cards (video/monitor, convergence, deflection and interconnect) and a multi-colour CRT; all are contained within a protective cover.
    • 38. The power supply units provide the requisite levels of a.c. and d.c. power necessary for overall operation of the display units.
    • 39. The video/monitor card contains a video control microprocessor, video amplifiers and monitoring logic for the display unit. The main tasks of the processor and associated EPROM and RAM, are to calculate gain factors for the three video amplifiers (red, blue and green), and perform input sensor and display unit monitor functions.
    14
  • 40. EFIS – Electronic Flight Instrument System
    Display unit (contd.)
    • The input/output interface functions for the processor are provided by analog multiplexers, an A/D converter and a multiplying D/A converter.
    • 41. The function of the convergence card is to take X and Y deflection signals and to develop drive signals for the three radial convergence (red. blue and green) coils and the one lateral convergence (blue) coil of the CRT.
    • 42. Signals for the X and Y beam deflections for stroke and raster writing are provided by the deflection amplifier card. The amplifiers for both beams each consist of a two-stage preamplifier, and a power amplifier. The amplifiers use two supply inputs, 15 V d.c. and 28 V d.c. the former is used for effecting most of the stroke writing, while the latter is used for repositioning and raster writing.
    15
  • 43. EFIS – Electronic Flight Instrument System
    Symbol Generators
    These provide the analog discrete and digital signal interfaces to the aircraft systems display units and control panel and they perform symbol generation system monitoring, power control and the main control functions of the EFIS.
    Remote Light Sensor
    This is a photodiode device which responds to flight deck ambient light conditions and automatically adjusts the brightness of the electronic displays to a compatible level.
    16
  • 44. EFIS – Electronic Flight Instrument System
    17
  • 45. EFIS – Electronic Flight Instrument System
    18
  • 46. EICAS – Engine Indicating and Crew Alerting System
    • The basic system comprises two display units, a control panel, and two computers supplied with analog and digital signals from engine and system sensors
    • 47. The system provides the flight crew with information on primary engine parameters (full-time) and with secondary engine parameters and warning / caution / advisory alert messages (as required).
    • 48. Display units
    The units are mounted one above the other, the upper unit displaying the primary engine parameters (EPR, N1 and EGT) and warning and caution messages, while the lower unit displays secondary engine parameters (N2, N3, fuel flow, oil quantity, pressure and temperature, and engine vibration), status of non-engine systems, aircraft configuration and maintenance data.
    19
  • 49. EICAS – Engine Indicating and Crew Alerting System
    • In the normal mode in flight, only the primary engine parameters are displayed; the lower display unit screen remaining blank.
    • 50. Warning, caution and advisory messages are displayed in red and yellow on the left-hand side of the upper display unit screen as conditions dictate. Abnormal secondary engine parameters are automatically displayed on the lower display unit.
    • 51. In each case the highest value attained is displayed in white under the actual readout, and the accumulated exceedance time is stored in a non-volatile memory of the computer for subsequent readout during maintenance mode checks.
    20
  • 52. EICAS – Engine Indicating and Crew Alerting System
    21
    Figure : EICAS Maintenance Panel
  • 53. EICAS – Engine Indicating and Crew Alerting System
    22
    Schematic functional diagram -EICAS
  • 54. EICAS – Engine Indicating and Crew Alerting System
    • During the normal mode of operation all secondary engine parameters may be displayed on the blank lower display unit by pressing an 'ENG' select switch on the EICAS control panel.
    • 55. A second switch (STATUS) is provided on the control panel and when pressed it switches the lower display into a mode that displays the status of several systems (e.g. flight control surface positions) and also up to 16 status messages requiring flight crew awareness prior to take-off and in flight.
    23
  • 56. ECAM – Electronic Centralized Aircraft Monitoring
    • The ECAM system comprises the units as far as display format is concerned, it differs significantly from EICAS in that it excludes analog presentation of engine parameters and it adopts the principle of mounting display units side-by-side so that the left-hand unit is dedicated to information on the system's status, warnings and corrective action in a sequenced checklist format, while the right-hand unit is dedicated to associated information in diagrammatic format.
    • 57. There are four display modes, three of which are automatically selected and referred to as flight phase-related, advisory (mode and status) and failure-related modes. The fourth mode is manual and permits the selection of diagrams related to any one of twelve of the aircraft's systems for routine checking. The selections are made by means of illuminated push-button switches on the ECAM control panel.
    24
  • 58. ECAM – Electronic Centralized Aircraft Monitoring
    25
  • 59. ECAM – Electronic Centralized Aircraft Monitoring
    • In normal operation, the automatic flight phase-related mode is used and in this case the displays are appropriate to the current phase of aircraft operation i.e. pre-flight, take-off, climb, cruise, descent, approach and after landing.
    • 60. The failure-related mode takes precedence over the two other automatic modes and the manual mode.
    • 61. STATUS messages, which are also displayed on the left-hand display unit, provide the flight crew with an operational summary of the aircraft's condition, possible downgrading of auto land capability, and as far as possible, and indications of the aircraft status following all failures except those that do not affect the flight.
    26
  • 62. ECAM – Electronic Centralized Aircraft Monitoring
    27
    Figure ECAM upper display
  • 63. ECAM – Electronic Centralized Aircraft Monitoring
    28
    Figure ECAM lower display
  • 64. ECAM – Electronic Centralized Aircraft Monitoring
    29
  • 65. ECAM – Electronic Centralized Aircraft Monitoring
    30
    Figure : ECAM Control Panel
  • 66. Flight Management System
    FMS comprises of -
    • Flight Management Computer (FMC)
    • 67. Two Control and Display Units (CDU)
    • 68. Thrust Mode Annunciator (TMA)
    • 69. Stored within the FMC is:
    Navigation Data Base - used to define route selection and contains airports, procedures, waypoints, navaidsetc. This portion of the internal data is inputted and updated by a portable data base loader and connector in the flight deck area.
    The CDU provides the interface between the crew and the FMC, and data exchange is provided by ARINC 429 busses.
    31
  • 70. Flight Management System
    32
    Figure : Flight Management System
  • 71. Flight Management System
    • Flight Management Computer
    Typically, a computer incorporates three different types of memory: a bubble memory for holding the bulk navigation and aircraft performance characteristics data bank; a C-MOS RAM for holding specific navigation and performance data, and the active and secondary flight plan, all 'down-loaded' from the bubble memory; and a UV-PROM for the operation program, which may be reprogrammed at card level.
    • The data base which is used for all computations contains numerous types of records in memory
    • 72. Any subsequent changes in navigation aids and procedures, and route structure changes, are also incorporated in the data base by means of the data loader, in accordance with a specified time schedule, e.g., a 28-day cycle.
    33
  • 73. Flight Management System
    34
  • 74. Flight Management System
    35
  • 75. Flight Management System
    Control Display Unit
    The CDU provides the primary means of flight profile
    selection/modification and display of associated parameters. The
    display format consists of 14 lines of data with a possible 24
    characters per line. The top line is the page title and number
    of pages associated with the display.
    • The bottom line is the "scratch pad" which is used for :
    Data entry.
    Message display.
    Transferring data field information.
    • Six lines of data is divided into right and left data fields with associated data titles.
    36
  • 76. Flight Management System
    • FMCS Messages
    These are displayed in the scratch pad line of the CDU. There are three categories of messages, and they have a defined priority for display should their set conditions occur at the same time. The three types listed in order of priority are:
    Alerting Messages
    Entry Error Advisory Messages
    Advisory Messages
    • Existence of a message in the scratch pad will illuminate the CDU's, Message (MSG) light and FMC Message will appear on the top CRT of EICAS. If the FMC fails then CDU 'FAIL' light illuminates and 'FMC Fail' appears on EICAS.
    37
  • 77. Flight Management System
    System Configuration
    • Two FMC systems are installed in an aircraft, each having its own CDU situated on the centre console and each controlling its associated automatic flight control system, auto-throttle system and radio-navigational aids. In normal operating conditions, both computers operate together and share and compare each other's information, i.e., they `cross talk' by means of an interconnecting data bus. The pilots can operate their displays independently for review or revision purposes without disturbance to the active flight plan and without affecting the other CDU commands.
    • 78. In the event of failure of one computer, each pilot has the means whereby he can select his own CDU into the other system.
    38
  • 79. Flight Management System
    39
    Figure : FMCS System Configuration
  • 80. Fly By Wire (FBW)
    • A considerable weight saving is achieved by replacing all the mechanical linkages (cables, rods, pulleys etc.) by electrical wires and computers.
    • 81. The A320 has a fly-by-wire system where the pilot's inputs are via
    a side stick positioned on his/her left side (the right-hand side for the
    second officer) and rudder pedals. The B777 is also fly-by-wire (with some
    fly-by-light) but has conventional flight deck controls (control wheels and
    rudder pedals).
    • With both aircraft, when the pilot puts an input into the flying control
    system, a signal from the control (in or near the flight deck) provides an
    electrical signal to the flying control computers. These provide electrical
    analogue signals to electrically controlled Power Flying Control Units
    (PFCUs) near the flying control surface.
    40
  • 82. Fly By Wire (FBW)
    • Conventional primary flight controls systems employ hydraulic actuators and control valves controlled by cables that are driven by the pilot controls. These cables run the length of the airframe from the cockpit area to the surfaces to be controlled. This type of system, while providing full airplane control over the entire flight regime, does have some distinct drawbacks.
    • 83. The cable-controlled system comes with a weight penalty due to the long cable runs, pulleys, brackets, and supports needed. The system requires periodic maintenance, such as lubrication and adjustments due to cable stretch over time.
    • 84. In addition, systems such as the yaw damper that provide enhanced control of the flight control surfaces require dedicated actuation, wiring, and electronic controllers. This adds to the overall system weight and increases the number of components in the system
    41
  • 85. Fly By Wire (FBW)
    • Control position and aircraft attitude is feed back to the computer so it can monitor the aircraft's movement.
    • 86. The main advantages of fly-by-wire are:
    • 87. Saving in weight.
    • 88. Requires less maintenance.
    • 89. More responsive.
    • 90. Increased fuel economy.
    • 91. Built in protection systems.
    • 92. Integration of several federated systems into a single system.
    • 93. Greater flexibility for including new functionality or changes after initial design and production.
    • 94. Ease of maintenance.
    42
  • 95. Fly By Wire (FBW)
    • A disadvantage is that the system is all electrical/
    electronic and electromagnetic interference such as lightning strikes could be a problem. The system has back-up facilities such a duplication and triplication of hardware and software. To overcome the possibility of complete electrical failure there is a mechanical standby mode. It gives limited flying control authority but if all else fails it can be used.
    43
  • 96. Fly By Wire (FBW)
    System Overview
    • In a FBW flight control system, the cable control of the primary flight control surfaces has been removed. Rather, the actuators are controlled electrically. At the heart of the FBW system are electronic computers. These computers convert electrical signals sent from position transducers attached to the pilot controls into commands that are transmitted to the actuators.
    44
  • 97. Fly By Wire (FBW)
    Because of these changes to the system, the following design features have been made possible:
    • Full-time surface control utilizing advanced control laws. The reaction time of the control laws is much faster than that of an alert pilot. Therefore, the size of the flight control surfaces could be made smaller than those required for a conventionally controlled airplane. This results in an overall reduction in the weight of the system.
    • 98. Integration of functions such as the yaw damper into the basic surface control.
    • 99. Improved system reliability and maintainability.
    45
  • 100. Fly By Wire (FBW)
    46
    Figure: The Primary Flight Control System on a typical aircraft is comprised of
    the outboard ailerons, flaps, elevator, rudder, horizontal stabilizer,
    and the spoiler/speed brakes.
  • 101. Fly By Wire (FBW)
    FBW System
    • An aircraft flight controls consist of primary and secondary systems. The primary systems - ailerons, roll spoilers, elevators, trimable horizontal stabilizer (tailplane) and rudder - control pitch, yaw and roll flight attitudes.
    • 102. The secondary systems comprise leading edge slats and trailing edge flaps for low speed flight handling, airbrakes/load alleviation spoilers for deceleration/load alleviation at all flight speeds and lift dumpers for deceleration after landing.
    • 103. An aircraft flight controls are hydraulically actuated and electrically or mechanically controlled.
    47
  • 104. Fly By Wire (FBW)
    FBW System
    • Input signals for the various controls are:
    Pitch Control
    Elevator Electrical
    Stabilizer Electrical for normal or alternate control. Mechanical for manual trim control
    Roll Control
    Ailerons Electrical
    Spoilers Electrical
    Yaw Control
    Rudder Mechanical (electrical for yaw damping, turn co-ordination and trim)
    48
  • 105. Fly By Wire (FBW)
    Load Alleviation Spoilers Electrical
    Slat/Flap Control Electrical
    Speed Brake ControlElectrical
    • All surfaces are hydraulically actuated.
    • 106. Controls in the cockpit consist of two side sticks, conventional rudder pedals and pedestal mounted controls and indicators.
    49
  • 107. Fly By Wire (FBW)
    AIRBUS FBW System
    • Electrical control is achieved by three types of computer: ELAC (Elevator and Aileron Computers) : These two computers achieve aileron control and normal elevator and stabiliser control.
    • 108. SEC (Spoilers and Elevator Computers) : These three computers achieve upper wing surfaces control, standby elevator and stabiliser control.
    • 109. FAC (Flight Augmentation Computers) : These two computers achieve rudder control.
    • 110. In addition the Flight Control Data Concentrator acquires data from the ELAC's and SEC's and sends this to the ECAM (Electronic Centralised Aircraft Monitor) - the flight deck screen displays and CFDS.
    • 111. The Electrical Flight Control System (EFCS) includes the ELACs, SECs, FCDCs and vertical accelerometers.
    50
  • 112. Fly By Wire (FBW)
    AIRBUS FBW System Arrangement
    51
  • 113. Fly By Wire (FBW)
    • AIRBUS FBW System
    • 114. The EFCS is designed according to the following principles:
    (a) Redundancy and dissimilarity.
    (b) Monitoring :
    • Cross-talk.
    • 115. Self-monitoring capacity.
    • 116. Monitoring channel: Each computer consists of two physically and electrically separated channels, one being dedicated to the control functions, the other to the monitoring functions.
    • 117. Automatic power-on and pressure-on safety tests performed without movement of the surfaces.
    52
  • 118. Fly By Wire (FBW)
    AIRBUS FBW System
    • Load Alleviation Function (LAP)
    -The load alleviation function, which operates through the ailerons and spoilers 4 and 5, becomes active only in condition of turbulence (pilot control authority is not modified) in order to relieve wing structure loads.
    -The high hydraulic demands required to achieve the rapid surface movements are provided with the help of dedicated hydraulic accumulators.
    • Yaw Control
    -The rudder is powered by three independent hydraulic jacks which operate in parallel and are controlled from the flight deck via a conventional cable flying control system.
    -Yaw mechanical control is by conventional manual rudder controls from the pilot's rudder pedals and is always available.
    53
  • 119. Fly By Wire (FBW)
    • Yaw Damping And Turn Co-ordination
    -In flight yaw damping and turn co-ordination functions are automatic.
    -In normal operation the three hydraulic servo jacks are driven by a green hydraulic servo actuator controlled by FAC 1. A yellow servo actuator controlled by FAC 2 remains synchronized and will take over in case of failure.
    -The yaw commands for turn co-ordination and yaw damping are computed by the ELACs and transmitted to the FACs. There is no feedback to the rudder pedals from yaw damping turn co-ordination functions.
    • Rudder Trim
    54
  • 120. Fly By Wire (FBW)
    The trim actuator and yaw damper servo actuators are used to introduce the A/P signals. The trim actuator drives the mechanical control (and pedals) through the artificial feel.
    55
  • 127. Fly By Wire (FBW)
    56
    Figure : Flight Control System - General Arrangement
  • 128. Fly By Wire (FBW)
    • In this mode the artificial feel breakout force is increased by a solenoid actuated device (100N - about 10kg - although kg is strictly a mass and not a force).
    • 129. The ELAC and SEC also feed two computers and the FCDC. These monitor and analyze ELAC and SEC maintenance messages at power on (on ground), in-flight and after touch down.
    • 130. It stores the data and delivers failure indications, eg stored failures, failure history, trouble shooting as well as failed LRUs to the Centralised Fault Display Interface Unit for onward transmission to the Maintenance Control DisplayUnit.
    57
  • 131. Fly By Wire (FBW)
    Boeing FBW System :
    • The Boeing 777, like the A320, has a highly integrated flying control system. Unlike the A320 it uses conventional flight-deck controls. Signaling is via ARINC 629 data buses, and various computers and control units. The control surfaces are hydraulically powered via PCUs.
    • 132. The flight deck controls consist of a control column for control of the elevators with a hand wheel for control of the ailerons, flaps, and roll spoilers. The rudder bar controls the rudder.
    • 133. These controls are provided with artificial feel and back-drive motors to move them in the correct sense when the system is in autopilot mode. An aileron trim actuator is also fitted in the system.
    58
  • 134. Fly By Wire (FBW)
    Boeing FBW System :
    • The ARINC 629 data bus is a twisted pair of wires transmitting data in both directions to all computers/LRUs (Line Replaceable Units). Each computer /LRU is connected to the bus by untwisting the twisted pair locally and clamping on an Inductive Couple Unit (which does not cut the insulation of the bus).
    • 135. Any computer/LRU can listen to any data on the bus and receive the data according to how its personality PROM (Programmable Read Only Memory) is programmed. In other words the computer's permanent memory knows what information on the data bus is for its use.
    59
  • 136. Fly By Wire (FBW)
    Boeing FBW System :
    • Flight deck control movement is converted into an electrical analogue signal by transducers (XDCRs) fitted to the flying control system under the flight deck floor. This signal is then sent to the Actuator Control Electronics LRU (ACE).
    • 137. The pilot's controls are connected via the Actuator Control Electronics (ACE) unit to the PCU. Other units such as the Primary Flight Computer (PFC) are connected into the system by the ARINC 629 bus.
    60
  • 138. Fly By Wire (FBW)
    Boeing FBW System :
    • The flight deck control is connected to position and force transducers, which signals the pilot's intention by an analogue signal to the ACE. This is in two way communication via the data bus with the PFC. An analogue command signal is sent to the PCU to move the ailerons in the desired direction. Positional feedback is sent to the ACE which controls the range and speed of movement of the PCU - and hence the control surface.
    • 139. The controls have artificial feel to simulate air loads on the control surfaces and trimming is achieved by biasing the system neutral by a trim unit actuator.
    61
  • 140. Fly By Wire (FBW)
    • Autopilot
    • 141. When autopilot is engaged the back-drive actuator will move the pilots controls in response to autopilot commands. Whenever the autopilot is engaged the back-drive actuators are active.
    • 142. When autopilot is selected the PCU is controlled by the ACE, PFC and Autopilot Flight Director Computer (AFDC) via the bus. The AFDC will also send an analogue signal to the back-drive actuator to move the flight deck controls to correspond to control surface movement. Thus the system simulates closely the characteristics of a conventional mechanical flying control system.
    62
  • 143. Fly By Wire (FBW)
    • As traffic demands increased, a need for more accurate navigation and positioning information became apparent. Two satellite navigation systems have been introduced to provide position and time to users around the globe. The global positioning system (GPS) provides accurate position and time information around the earth, 24 hours a day.
    • 144. The Global Positioning System (GPS) is a satellite navigation system providing accurate, three-dimensional position, velocity, and time information all over the world. GPS is a network of satellites that broadcasts ranging information, in addition to the satellite position, and time of transmission.
    63
  • 145. Fly By Wire (FBW)
    • The increased accuracy of an aircraft's position using GPS provides a number of navigation benefits. These benefits would include:
    (1) reduced separation of aircraft without compromising safety,
    (2) optimizing aircraft routes,
    (3) improved approach/departure and
    (4) improved traffic control. All of these benefits will reduce cost and increase safety of air travel.
    • The purpose of the Ground Proximity Warning System (GPWS) is to alert the flight crew to the existence of an unsafe condition due to terrain proximity, or a windshear condition.
    64
  • 146. Global Positioning System :
    65
  • 164. Global Positioning System :
    • GPS-Space Segment:
    Space Segment 1- General Description
    • 24 Satellites
    -21 in use
    -3 active spares
    • 6 Different orbital paths
    -4 satellites in each
    • 1 Complete orbit in 12hrs
    -Orbit altitude 10898 Nm (20183km)
    • Each orbital planes is included 55º and contains four satellites.
    • 165. 6 - 11 Satellites in view at 5 degrees or more above the horizon to users, anytime and anywhere in the world.
    66
  • 166. Ground Positioning System :
    • GPS-Space Segment:
    Space Segment 2- General Description
    Satellite frequencies:
    • Frequency band,L1 - 1575.42 MHz: Navigational Message, (C/A) code, P(Y)[Precision] code
    • 167. Frequency band,L2 - 1227.60 MHz: P(Y) code
    • 168. The satellites transmit at two L-band frequencies to allow the detection of signal propagation delay in the ionosphere.
    • 169. Each satellite's P-code and C/A code is a unique PRN code. The unique code is a pseudo-random sequence that allows all the satellites to transmit on the same frequency without creating radio interference.
    67
  • 170. Ground Positioning System :
    Space Segment 3 – GPS Transmission
    • GPS Data comprises of:
    • 171. Universal Time – Co-ordinated UTC
    • 172. Clock Corrections
    • 173. Ephemeris Parameters
    • 174. Orbital data giving satellite positions
    • 175. Almanac data
    • 176. Satellite housekeeping data
    • 177. Transmitted on L band frequencies
    • 178. L1 – 1.57542 GHz
    • 179. L2 – 1.227 GHz
    • 180. Transmissions are spectrum spread
    • 181. All satellites use the same frequencies
    68
  • 182. GlobalPositioning System :
    69
    Space Segment 3 – GPS Transmission
  • 183. GlobalPositioning System :
    Control Segment 1 – General Description
    • Controls satellites
    Master station
    Five Monitoring station (receiver-transmitters, antennas)
    - Allow all satellites to be seen at any one time
    70
  • 184. GlobalPositioning System :
    Control Segment 2 – General Description
    • The visible satellites and monitor the broadcast information is sent to a master control station (located in Colorado)that computes the satellite orbits (ephemeris data), and satellite clock corrections for each satellite. Three of the ground stations have an uplink capability.
    • 185. Control Segment 3 – Master Control Station
    • 186. Located at Colorado Springs, USA
    • 187. Controls all the control segment
    • 188. Equipped with atomic clock
    - Reference for GPS satellites
    • Uses upload monitoring stations to send:
    - Orbit correction commands
    - Navigation messages
    71
  • 189. GlobalPositioning System :
    • Control Segment 3 – Master Control Station
    72
    Master Control Station
    Monitoring/Upload Stations
  • 190. GlobalPositioning System :
    • CONTROL SEGMENT 4 – MONITORING STATIONS
    • 191. 5 Monitoring Stations
    - Hawaii
    - Cape Canaveral
    - Ascension Island
    - Diego Garcia Island
    - KawajaleinIsland
    • 3 Uploading Station
    • 192. Receive information from satellites
    - Accuracy of satellite clocks
    - Meteorological data
    - To calculate tropospheric signal delay
    - Measure ranges of all visible satellites
    - To calculate and predict orbits
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  • 193. GlobalPositioning System :
    • Control Segment 4 – Monitoring Stations
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  • 194. GlobalPositioning System :
    • User Segment
    • 195. GPS Receiver
    - Receives GPS Data
    • Data derived by receiver
    -Longitude, latitude and height
    -Calculated by GPS receiver
    -Ground speed
    -Actual track
    -Position error
    • The GPS may be coupled with other systems (navigation data bases, ACARS, ATC, etc) or sensors (radio altimeters, baro altimeters; DMEs, etc) to provide increased situational awareness enrouteand more runway detail during approaches/departures.
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  • 196. GlobalPositioning System :
    User Segment 2 – GPS Receiver
    Decodes GPS Signal
    3 Types used on aircraft
    • Multi Mode Receiver (MMR)
    Located in MEC
    • GPS Signal Unit (GPSSU)
    Stand alone
    Located near GPS antenna
    • GPS Sensor Module
    On circuit board inside LRU
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  • 197. GlobalPositioning System :
    GPS ERRORS
    The range measurement is the basis of the receiver determining the location of the receiver. There are several other contributing errors that still are capable of reducing the accuracy of the range measurement. These range errors are:
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  • 203. Global Positioning System :
    • Uses :
    1.Allows Calculation ofof
    2.Assist aircraft T/O, landing & taxiing
    3. Gives 3 dimensional positioning
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  • 207. Inertial Reference System:
    • The Inertial Reference System (IRS) and Inertial Navigation System (INS) are methods of very accurate navigation that do not require any external input such as ground radio information. They are passive systems that work entirely independently of any external input. The Inertial Reference Systems (IRS) which mainly provide: Attitude, Heading, Ground Speed and Present Position information.
    • 208. Components of IRS
    The complete IRS consists of:
    IRU (Inertial Reference Unit)
    MSU (Mode Selector Unit)
    ISDU (Inertial System Display Unit)
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  • 209. Inertial Reference System:
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  • 210. Inertial Reference System:
    • Components of IRS
    • 211. The ISDU enables the IRS to be initialized and their main navigation data to be displayed. The unit is an emergency means for these operations which are normally performed through the FMS.
    • 212. The MSU allows the related IRS to be started up and the operation mode selection. Warning lights enables the system monitoring.
    • 213. The IRU is the main component of this strap down inertial system. It houses 3 (three) accelerometers and 3 (three) ring laser gyros which detects the aircraft rotations and accelerations around its 3 (three) axes, and the computer which calculates attitude, direction, speed and reference position.
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  • 214. Inertial Reference System:
    • Mode Selector Unit (MSU)
    The mode select unit (MSU) is designed to interface with one IRU. It is used to command the modes of operation with the mode selector knob:
    • OFF MODE: All the circuitry is de-energized.
    • 215. NAV MODE: Normal position for IRS operation. After alignment phase, the IRS will provide attitude, heading and position information to all peripherals
    • 216. ATT MODE: Selection of this mode when the alignment is not complete.
    Only attitude and heading information is usable. The MSU is also used to annunciate status of the IRU.
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  • 217. Inertial Reference System:
    • Inertial Reference Unit (IRU)
    The IRU is the system main unit. Outside the gyros and accelerometers, the IRU contains power supplies and eight cards of electronics for sensor signal conditioning, data handling, timing and control, data processing, and system interface.
    Each IRU consist of the following module:
    Sensor assembly
    PCB modules
    Rear Panel Assembly
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  • 218. Inertial Reference System:
    • Inertial Sensor Display Unit (ISDU)
    • 219. The ISDU is connected to the IRU and allows their initialization. It is also used to display the main navigation data such as present position, ground speed, true heading, wind information and system status. It is designed to interface with discrete and digital inputs received from each IRU.
    • 220. It can transmit discrete signals or digital messages to the IRU. The ISDU consists of a front-panel mounted display, two selector knobs: SYS DSPL and DSPL SEL and a keyboard.
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  • 221. Inertial Reference System:
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    Figure: Inertial Sensor Display Unit (ISDU)
  • 222. Traffic Alert And Collision Avoidance System (TCAS)
    • TCAS is a family of airborne devices that function independently of the ground-based air traffic control (ATC) system and provide collision avoidance protection for a broad spectrum of aircraft types.
    • 223. TCAS I provides traffic advisories (TA) and proximity warning of nearby traffic to assist the pilot in the visual acquisition of intruder aircraft.
    • 224. TCAS II provides traffic advisories and resolution advisories (RA), i.e., recommended escape maneuvers, in the vertical dimension to either increase or maintain the existing vertical separation between aircraft.
    • 225. TCAS transmits to and receives signals from other airplanes to get altitude, range, and bearing data.
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  • 226. Traffic Alert And Collision Avoidance System (TCAS)
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    Figure : TCAS airspeed coverage
  • 227. Traffic Alert And Collision Avoidance System (TCAS)
    TCAS Components
    These are the TCAS components:
    • TCAS directional antennas
    • 228. TCAS computer
    • 229. Air traffic control (ATC)/TCAS control panel.
    ATC Control Panel
    • A single control panel is provided for ATC/TCAS to allow the flight crew to select and control all TCAS equipment, including the TCAS Processor, the Mode S transponder, and in some cases, the TCAS displays.
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  • 230. Traffic Alert And Collision Avoidance System (TCAS)
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    Figure TCAS Basic Operation
  • 231. Traffic Alert And Collision Avoidance System (TCAS)
    TCAS Display/Indication
    The traffic collision avoidance system (TCAS) computer puts traffic into these four groups:
    • Other traffic shows as a white open diamond; the altitude readout is in white text
    • 232. Proximate traffic shows as a solid white diamond; the altitude readout is in white text
    • 233. Traffic advisory (TA) shows as a solid amber circle; the altitude readout is in amber text
    • 234. Resolution advisory (RA) shows as a solid red square; the altitude readout is in red text.
    • 235. Each traffic symbol has an altitude readout. A vertical motion arrow is also shown if the airplane vertical speed is greater than 500 feet per minute (fpm).
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