M13.02- Communication - Navigation- ATA 34

For Training Purpose Only DETAILED TRAINING VAR Part 7 - Aircraft Maintenance Basic Cat B2 TRAINING MANUAL ATA 34 Issue: 01 Rev: 00 Date: 25/04/2014 © VAECO Training Center Part-66 ATA 34 NAVIGATION For Training Purposes Only Lufthansa Technical Training M11.5.5 / M13.4 NAVIGATION ABBREVIATIONS HAM US/F-4 SaR Dec 2005 Page 1 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 ABBREVIATIONS A α α-floor AAAS A/C A/C ACARS ACC ACT ADC ADF ADI ADIRS ADIRU ADM ADS ADS AFS AGL AHRS AIMS AIP alpha ALT AOA A/P APL Alpha angle of attack max. AOA, stall Alternate Audio Alert Select air conditioning aircraft ARINC/aircraft communication, addressing and reporting syst. automatic clearance control actual air data computer automatic direction finder attitude director indicator air data and inertial reference system air data and inertial reference system unit air data module Air Data System automatic dependant surveillance autoflight system Above Ground Level attitude and heading reference system Airplane Information Management System aeronautical information publication angle of attack altitude angle of attack autopilot airplane HAM US/F-4 SaR Dec 2005 appr APU ARR ARINC ARINC 407 ARINC 419 ARINC 423 ARINC 429 HS ARINC 429 LS ARINC 453 ARINC 568 ARINC 575 ARINC 591 ARINC 604 ARINC 629 ARINC 706 ARINC 717 ARINC 741 ASP at atm A/T ATC ATN ARINC ASL ATP ATT approach auxillary power unit arrival Aeronautical Radio Incorporation (non profit organization) analog synchro data transmission standard for ADC altitude digital data transmission standard for civil A/C (7.5-14.5 kBaud) report for standardization of design and use of BITE digital data transmission standard for civil A/C (100kBaud) digital data transmission standard for civil A/C (12-14.5 kBaud) digital data transmission standard for WXR (1MBaud); EGPWS digital data transmission standard for DME in A/C digital data transmission standard for ADC altitude (using 419) digital data transmission standard between DMU and DAR incorporates ARINC 423 digital data transmission standard for B777 digital data transmission standard for ADC altitude (using 429) digital data transmission standard between DFDAU and DFDR aviation satellite communication system audio selector panel atmosphere, 1at = 0.980665 bar atmosphere, 1atm = 1.01325 bar autothrottle air traffic control (transponder) aeronautical telecommunications network Aeronautical Radio Incorporation Above Sea Level Acceptance Test Procedure attitude Page 2 Part-66 B baud B/B BCD BFE birdy BIST BIT BITE BNR BRT Btu Bravo bits per second back beam (approach) Binary Coded Decimal buyer furbished equipment indication for maximum angle of attack Built in Self Test Built In Test built in test equipment Binary bright 1Btu = 1.05506 kJ C C CAA CAIMS cal CAPT CAUT CAS CAT CDU CFDS Charly degrees Celsius Civil Aviation Authority Central Aircraft Information/Maintenance System calorie, 1cal = 4.1868 J captain caution corrected/calibrated/computed airspeed category (of weather for landing) control and display unit centralized fault and display system For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS HAM US/F-4 SaR Dec 2005 CFIT CFM CG CL CLB CLR CIT CMC CMD CMC C/O COMP CO-route CP CP crab angle CRS CRZ CRT CT cu ft cu in CW CCW CWS Controlled Flight Into Terrain Cubic Feet per Minute center of gravity climb climb clear compressor inlet temperature centralized maintenance computer command centralized maintenance computer Callouts computer company route Control Panel center of pressure angle between track and heading (VOR/LOC) course cruise cathode ray tube control transformer (synchro) cubic feet, 1cu ft = 0.028317 m3 cubic inch, 1cu in = 16387,064 mm3 Clock Wise Counter Clock Wise control wheel steering Page 3 Part-66 D Delta difference, error signal PSC DAA DADC DAU DDM DEP DEV DFCS DG DH DISPL DITS DME DN DO drift angle DSR TK DSP DSU DSWC preselected course error digital analog adapter Digital Air Data Computer Data Acquisition Unit Difference in Depth of Modulation departure deviation digital flight control system directional gyro decision height display Digital Information Transfer System distance measurement equipment down Discrete Output angle between heading and track desired track Digital Signal Processor Display Switching Unit Digital Stall Warning Computer E EADI EAS ECAM ECU EEC EEPROM EFIS EFCP EGPWC EGPWD EGPWS EGT EHSI EICAS EIS EIVMU EMER EMI ENB EPR EPROM Echo electronic attitude director indicator - PFD equivalent airspeed, engineering value, not indicated in cockpit electronic centralized aircraft monitoring engine control unit electronic engine control Electrically Erasable Programmable Read Only Memory Electronic Flight Instrument System EFIS Control Panel Enhanced Ground Proximity Warning Computer Enhanced Ground Proximity Warning Display Enhanced Ground Proximity Warning System exhaust gas temperature electronic horizontal situation indicator - ND Engine Indication and Crew Alert System electronic/engine instrument system engine indication and vibration monitoring unit emergency Electromagnetic Interference Enabled engine pressure ratio Erasable Programmable Read Only Memory For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS HAM US/F-4 SaR Dec 2005 Page 4 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 F FAA FADEC FCC FCU FCU F/D FDR FET FF FIAS Fix flt flag FMC FMCS FMGC FMGEC FMS FMU FO FPM FSEU ft FWC FWD F/T Foxtrott Federal Aviation Administration full authority digital engine control Flight Control Computer flight control unit fuel control unit flight director Flight Data Recorder field effect transistor fuel flow Flight Inspection Aircraft System approach points (initial-, intermediate-, and final-approach-fix) flight warning indication for failed system Flight Management Computer flight management computer system flight management and guidance computer flight management guidance and flight envelope computer Flight Management System fuel metering unit first officer Feet per Minute Flaps/Slats Electronic Unit feet, 1 foot = 0.3048 m Fault Warning Computer forward Functional TestF/W Fail/Warning G g Gal GDOP GEN Golf 9.80665 m/sec2 gallon, see ”US Gal” and ”Imp Gal” Geometric Dilution of Position generator HAM US/F-4 SaR Dec 2005 GLS GMT gnd GPS GPWS GS G/S GPS Landing System Greenwich Mean Time ground (terrestrial or electrical) Global Position System Ground Proximity Warning System ground speed Glideslope H h hardover hdg HDOP HI hld hldg HMU hp HP hPa HPT HS 429 HSI HSID H/W Hotel hour undesired sudden movement of A/C by faulty A/P signal heading Horizontal Dilution of Position high hold holding hydro mechanical unit horse power, 1hp = 0.7457 kW high pressure (stage) hecto Pascal: 100 Newton per square meter high pressure turbine ARINC standard 429, high speed, 100 kbaud horizontal situation indicator Hardware/Software Interface Document Hardware Page 5 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 I IAS ident IDU IFR ILS IM IMC Imp Gal in in Hg IND INIT INOP INS I/O IOC IRS ISA ISO IVS India indicated air speed, necessary for aerodynamics special identification for ATC indication display unit instrument flight rules Instrument Landing System inner marker instrument meteorological conditions Imperial Gallon, 1 Imp Gal = 4.546 liter inch, 1in = 25.4 mm inches mercury, 1in Hg = 33.8639 mbar indicator inital (FMS page) Inoperative inertial navigation system Input/Output Input/Output Concentrator inertial reference system international standard atmosphere International Standards Organization Inertial Vertical Speed J Juliett K K kcal kcal/h km knot Kilo Kelvin kilocalorie, 1kcal = 4.1868 kJ power, 1kcal/h = 1.1630 W Kilometer 1 NM per hour = 1.852 Km per hour HAM US/F-4 SaR Dec 2005 kp kts kilopond, 1kilopond = 9.80665 N knots L l L LAT Lb Lbf LCD LCTR LDG LED LG LH LH LNAV LO LOC LON LP LPT LRRA LRU LS LS 429 LSK LSB LT LVDT LVL CHG Lima liter left lattitude pound, 1lb = 0.45359237 kg pound force, 1lbf = 4.44822 N liquid cristal display locator landing Light Emitting Diode landing gear left hand Lufthansa lateral navigation low localizer longitude low pressure (stage) low pressure turbine Low Range Radio Altimeter Line Replaceable Unit Landing System ARINC standard 429, low speed, 12-14.5 kbaud line select key Least Significant Bit light linear variable differential transducer level change Page 6 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 M m max min min mm mph MAC magn MASI max MCDU MCP MCT MDA MEC MEL MFD MIC MKR MK II / V / VII MLS MM MMO MMR mode A mode B Mike meter maximum minimum minute millimeter (statute) miles per hour, 1mph = 1.609344 Km/h mean aerodynamic cord magnetic Mach airspeed indicator maximum multipurpose control and display unit Mode Control Panel maximum continous thrust minimum decent altitude main engine control minimum equipment list Multi−Functional Display microphone marker Mark Two / Five / Seven Warning Computer microwave landing system middle marker Mach maximum operating Multi Mode Receiver transponder mode A: 4 digits octal code transmission transponder mode B HAM US/F-4 SaR Dec 2005 mode C mode D mode S MOPS MOS MSB MLS MSL MTBF transponder mode C: altitude reporting transponder mode D transponder mode S: selective, necessary for TCAS minimum operational performance standards (for TWDL) metal oxide semiconductor Most Significant Bit Microwave Landing System mean sea level Mean Time Between Failure N N/A NCD N1 N2 ND NDB NIL NM NO NVM November Not Applicable No Computed Data number of cycles per second of the fan shaft number of cycles per second of the core engine shaft navigation display - EHSI non directional beacon not in list 1 nautical mile = 1.852 km normal operation Non Volatile Memory O OPS ONS OM OMS oz Oscar operations omega navigation system outer marker onboard maintenance system ounze, 1oz(force) = 0.278014 N, 1oz(mass) = 28.349523 g Page 7 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 P PAPI PAD (bit) PAR P/B PC PCMCIA PDC PFD PMC PMAT P/N PNL POS PP PPI PS PSC psi PTT PVM PWR PWS Papa precision approach position indicator unused bit Parity push button Personal Computer Personal Computer Memory Card Industry Association performance data computer Primary Flight Display - EADI power management computer Portable Maintanence Access Terminal Part Number panel position Program Pin Plan Position Indicator horse power(german), 1PS = 0.73549875 kW preselected course pound per square inch: 14.5 psi = 1 bar, 1psi = 68.9476 mbar push to talk Processor/Voice/Memory power (electric, engine, hydraulic, pneumatic...) Predictive Windshear System Q QDM Quebec magnetic track to the station HAM US/F-4 SaR Dec 2005 QDR QFE QFU QNE QNH QTE QTY QUJ magnetic bearing from the station atmosheric pressure at airport elevation runway heading altimeter indication related to standard pressure of 1013.25 hP pressure at airport calculated to mean sea level true bearing from the station quantity true bearing to the station R R RA RA RAM rcdr rcvr RDDMI RH RMI RNAV ROM RPM RTCA RTE RVDT RWY RX Romeo right resolution advisory (TCAS) radio height Random Access Memory recorder receiver radio direction distance magnetic indicator right hand radio magnetic indicator area navigation Read Only Memory rotations per minute Requirements and Technical Concepts for Aviation route rotary variable differential transducer runway torque receiver synchro Page 8 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 S s sec SAT SDAC SDAC SDI SEL SFCC SFE SFFS SID SIG SPC SPD SPI sp ft sq in SSM S/T stall STAR statute mile STCA STD SW S/W SYS Sierra second second static air temperature system data digital analog converter (A310) system data acquisittion concentrator (A320) source destination identifier select(or) slat flap control computer seller furbished equipment System Flight Fault Summary standard instrument departure Significant Stall Protection Computer speed additional pulse for ”ident” (ATC) square feet, 1 sq ft = 0.092903 m2 square inch, 1 sq in = 645.16 mm2 sign status matrix Self Test breakdown of lift, IAS/CAS below minimum speed standard terminal arrival route 1 statue mile = 1.609344 km short term conflict alert, conflict warning at ATC ground station standard (atmospheric pressure, 1013,25 hPa) switch Software system T t t T Tango temperature, measured in Celsius time absolute temperature, measured in Kelvin HAM US/F-4 SaR Dec 2005 TA TACAN TAD TA&D TAS TAT TBD TA&D TCAS TCF TIT TK TKE T/O TOGA TORA Torr TSO TWDL TWR TX traffic alert (TCAS) Tactical Air Navigation Terrain Awareness Display Terrain Awareness & Display true airspeed, necessary for navigation total air temperature to be determined Terrain Awareness & Display Traffic Collision Avoidance System Terrain Clearance Floor turbine inlet temperature track track angle error take off take off go around (switch) take off range available pressure, 1Torr = 1mmHg = 1.33322 mbar Technical Standing Order two-way data link control tower torque transmitter synchro Page 9 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 U UART ULB US Gal USM UTC UUT Uniform Universal Asynchronous Receiver Transmitter under water locator beacon United States gallon, 1 US Gal = 3.785 liter Unsigned Magnitude universal time coordinated = replaced former GMT Unit Under Test X Xceiver X-feed X-fer X-mitter X-ponder XTK X-ray transceiver crossfeed transfer transmitter transponder cross track V VASIS VBV VDC VDOP VFR VG VHF VLSI VMC VMO VNAV VOR VORTAC V/S VSI VSV V-bars Victor visual approach slope indicator system variable bleed valves Volts, DC Vertical Dilution of Precision visual flight rules vertical gyro very high frequency Very Large Scale Integrated Circuit visual meteorological conditions velocity maximum operating vertical navigation VHF omnidirectional range VOR-TACAN vertical speed vertical speed indicator variable stator vanes type of flight director indicator pointers Y yd Yankee yard, 1 yd = 0.9144 m Z zulu Zulu zulu time = UTC W warn waypoint WC WP W/S Whisky warning enroute navigation fix Warning Computer waypoint Windshear HAM US/F-4 SaR Dec 2005 Page 10 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ABBREVIATIONS Part-66 Basic Value Basic Unit Length Mass Time Current Temperature Molar Magnitude Candle Power Symbol Prefix Symbol Meter Kilogram Second Ampere Kelvin Mole Candela Value m kg s A K mol cd Power of Ten For Training Purposes Only engl, Europe / US,France T G M k h da d c m n p f a Tera Giga Mega Kilo Hecto Deca Deci Centi Milli Micro Nano Pico Femto Atto HAM US/F-4 SaR Billion/Trillion Millard/Billion Million Thousand Hundred Ten Tenth Hundredth Thousandth Millionth Millardth/Billonth Billonth/Trillonth ...../Quadrillionth Trillonth/ Quintillonth Dec 2005 1012 109 106 103 102 101 10-1 10-2 10-3 10-6 10-9 10-12 10-15 10-18 Variable Length Area Volume Unit meter squaremeter cubic meter Liter Mass gram kilogram ton Time second minute hour Force Newton Energy Joule Kilowatt-hour Power Watt Kilowatt Torque Newton-meter Pressure Pascal and hecto Pascal Tension Bar Current Ampere Temperature Kelvin Degrees Celsius Viscosity dynamic Pascal-second kinematic squaremeter per second Symbol Relationships m m2 m3 l g kg t s min h N J kWh W kW Nm P hP bar A K C Basic Unit 1m2 = 1m * 1m 1m3 = 1m * 1m * 1m 1l = 0.001m3 1g = 0.001kg Basic Unit 1t =1000 kg Basic Unit 1min = 60 s 1h = 3600 s 1N = 1kgm/s2 1J = 1Nm = 1Ws 1kWh = 3.6*106Ws 1W = 1Nm/s = 1J/s 1kW = 1000W Pa s m2/s 1Pa =1Ns/m2 1Pa = 1N/m2 1hPa = 1mbar 1bar = 105Pa Basic Unit Basic Unit Kelvin minus 273.15 Page 11 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION VOR Part-66 VOR GENERAL The VOR ( VHF Omnidirectional Radio Range ) system is a Radio Navigation Aid used for short - and medium Range. Since 1949 the VOR is a recommended and standardized navigation aid by the ICAO. The following VOR navigation procedures are possible: S Area navigation S Holding S VORDME Approaches The function of the VOR system is to receive, decode and process bearing information transmitted from the VOR groundstation-signal. Mainly the VOR is used for area navigation. The groundstation are transmitting courses - so−called RADIALS- which are referenced to magnetic north. On these RADIALS the aircraft flies from one groundstation to the next. For the position determination of the A/C, two VOR groundstation are necessary. The A/C position is given by the intersection of two radials ( Theta/Theta- Navigation ). In conjunction with the DME ( Distance Measuring Equipment ) and a Navigation Computer on board, nonradial courses can be flown. This - also ICAO standardized procedure - can be used for areanavigation instead of navigation linked to airways. VOR gives the direction and DME the distance ( Theta/Rho- Navigation ). Instead of DME also TACAN ( Tactical Air Navigation ) and VOR ( VORTAC ) can be used for areanavigation. VOR stations, which are installed in the vicinity of airports, and which are used mainly for holdings or VORDME Approaches, are called Terminal VOR (TVOR). Beside the main navigation purpose, the VOR groundstation is used to transmit traffic information e.g. weather. For identification each VOR groundstation transmits a three letter Morse code. HAM US/F-4 SaR Dec 2005 Frequ. Range : Channel Modulation : : Range Tuning : : 108.00 - 117.95 Mc 108,00 - 111,90 Mc TVOR only at even tenth Mc 50 Kc 30 Hz FM reference phase , 30 Hz AM variable phase , 1020 Hz identifier, 300-3000 Hz voice Approx. 50 - 200 Nautical Miles S Manual at a VHF NAV control panel, S Automatically by FMGECs or S Manually by the MCDU’s or S by RMPs in NAV BACK UP MODE Page 12 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 For Training Purposes Only e.g. Figure 1 HAM US/F-4 SaR Dec 2005 VOR General Page 13 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 VOR charts For enroute, arrival and approach different charts exist. VOR ground stations Different types of ground stations are in use for VOR operations. They differ in their combination with other navigation aids. Another difference may be the transmitter power: high power for enroute navigation and low power for terminal VORs. Possible VOR stations are: S S S S S S VOR DVOR VOR DME VOR TACAN TVOR VOT Doppler VOR Terminal VOR Test VOR For Training Purposes Only Frequency information are in the approach charts and in the enroute maps. The three letters for each station are the Morse code identifier and the numbers are the frequency in MHz. The ”D” in front of the frequency indicates, that a DME station is collocated with the VOR station. VORs with high power transmissions of about 200 W, are for enroute navigation. The range should be more than 100 NM. Terminal VORs are transmitting only with about 25 W. The range is between 25 and 50 NM. Sometimes, these stations transmit ATIS information additionally to the identification code. Figure 2 HAM US/F-4 SaR Dec 2005 VOR Approach Chart Page 14 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Figure 3 HAM US/F-4 SaR Dec 2005 VOR Arrival ( STAR ) Chart ( Jeppesen ) Page 15 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 BEARING DEFINITION Bearing is the measurement of an angle between a reference line and a line between aircraft and ground station. Due to the selected reference line are three different bearings: S Magnetic Bearing MB S Relative Bearing RB S ( True Bearing TB ) Note: The True Bearing is not used in conjunction with VOR, so it is not described in this section. Mathematical relations between the terms are: QDM = RB + MH RB QDR = QDM = QDM + MH 1800 For Training Purposes Only Magnetic Bearing ( see Figure 4 A ) The magnetic bearing is a bearing, that is referenced to magnetic north. The VOR ground station are aligned to magnetic north and so the courses are magnetic bearings. The magnetic bearing from a VOR station is called a RADIAL. From the time of Morse- telegraphy, the bearings are also expressed as Q-codes. The magnetic bearing from the ground station to the aircraft is called QDR. The magnetic bearing from the aircraft to the ground station is called QDM. The relation between QDM and QDR is always 180 Relative Bearing ( see Figure 4 B ) The relative bearing is a bearing, that is referenced to the aircraft lubber line. To get a magnetic bearing out of a relative bearing - e.g. received from ADF - the RB has to be linked to the magnetic heading ( MH ) from a compass system. HAM US/F-4 SaR Dec 2005 Page 16 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 MN MN RADIAL A: Magnetic Bearing QDM MN MN For Training Purposes Only RADIAL B: Relative Bearing QDM RB Figure 4 HAM US/F-4 SaR Dec 2005 Magnetic- / Relative- Bearing Page 17 Part-66 VOR PRINCIPLE The principle of the VOR function can be easily illustrated by a lighthouse. A lighthouse, surrounded by water, is equipped with a focused rotating white beam, that rotates with 1 /sec. Every time the beam is exactly directed to magnetic north, a red flash light is switched on, which can be seen from any direction. If a sailor measures the time , passed from the flash of the red light, till the presence of the white beam, the time represents the angle or bearing between magnetic north and the line of position ( LOP ) of the ship. The LOP is called radial. To determine the radial, two signals are necessary: S a reference signal ( red flash light ) S a variable signal ( white rotating beam ), which is variable with the position of the receiver The timedifference corresponds to the bearing information. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR HAM US/F-4 SaR Dec 2005 Page 18 Part-66 Red flashlight White rotating Beam For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Figure 5 HAM US/F-4 SaR Dec 2005 Lighthouse principle Page 19 Part-66 VOR GROUNDSTATION There are two types of VOR groundstation: S Conventional VOR S Doppler VOR Conventional VOR Instead of light, the VOR uses radiosignals. The rf carrier is modulated with two 30 Hz signals. The phase of one 30 Hz signal is independent from position of the receiver. The phase of this 30 Hz signal is the Reference Phase. This phase corresponds to the red light. The second 30 Hz signal is transmitted in a way, that its phase varies with the azimuth.The phase of this 30 Hz signal is the Variable Phase and corresponds to the white light beam. The Phasedifference corresponds to the bearing information. Figure 6 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR HAM US/F-4 SaR Dec 2005 VOR conventional ground station Page 20 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Figure 7 HAM US/F-4 SaR Dec 2005 Conventional VOR Page 21 Part-66 Frequency spectrum of VOR The VOR station transmits a reference and a variable phase of 30 Hz. S Reference Phase The 30 Hz reference phase is FM modulated on a subcarrier of 9960 Hz with a frequency variation of 480 Hz. This subcarrier is modulated as AM on the selected carrier out of 108 to 117.95 MHz. This signal is transmitted via the omnidirectional antenna. S Variable Phase The variable phase consists out of the carrier frequency only, transmitted by a rotating dipole, overlayed by an omnidirectional signal. The rotating dipol pattern and the omnidirectional pattern result in the air in a rotating cardioide antenna characteristic. The cardioide makes 30 turns in one second and so the result is an 30 Hz amplitude modulated ( AM ) carrier signal in the receiver. Additional modulation is S two or three letters Morse code identifier ( IDENT ) with 1020 Hz and S somtimes voice information ( ATIS). In the VOR receiver, reference and variable phase are seperated and the phase difference determines the radial, on which the aircraft is located. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION VOR HAM US/F-4 SaR Dec 2005 Page 22 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 FIELD OF THE OMNIDIRECTIONAL ANTENNA RESULTANT FIELD FIELD OF THE DIPOLE ANTENNA VOR Antenna Pattern CARRIER FREQUENCY VARIATION 480 Hz FREQUENCY VARIATION 480 Hz VOICE VOICE fC For Training Purposes Only fC - 1020 Hz IDENTIFIER fC - 9960 Hz fC - 30 Hz 30 HZ FM Variable Figure 8 Dec 2005 fC + 30 Hz fC + 9960 Hz 30 HZ FM 30 HZ AM Reference HAM US/F-4 SaR fC + 1020 Hz IDENTIFIER Reference Conventional VOR Spectrum Page 23 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION VOR Part-66 Doppler VOR The DVOR uses the Doppler effect to generate a compatible VOR signals. In some cases, e.g. mountainous areas, a conventional VOR may lead to misnavigation due to reflected signals. The solution is a (more expensive) DVOR. Decisive for the better performance is the exchange function of the the two 30 Hz modulations. S Reference Phase The 30 Hz reference phase is now AM modulated on the carrier and is transmitted omnidirectional via the center ntenna. S Variable Phase The 30 Hz varible phase is now as a FM modulation on the subcarrier of 9960 Hz. As a result of the FM, if a high modulationindex is used, the variable phase is less interference−prone compared with the variable AM of the conventional VOR. The Transmission of the groundsignals is done in a way, that the exchange of reference - and variable - phase has no effect on the indication. Generation of the variable phase The variable phase is generated by switching the carrier signal - modulated by 9960 Hz AM - sequentially in a ring of dipoles from one dipole to the next. 30 full cycles in a second. From the view of the VOR receiver, sometimes the signal seems to move towards the aircraft (this causes an increase in frequency due to the Doppler effect), and after passing the signal moves away from the aircraft (a decrease of the frequency is the result). The 9960 Hz are now automatically frequency modulated with 30 Hz. The frequency variation of 480 Hz ar achieved by the diameter of 13.5 m of the ring of antennas. Figure 9 HAM US/F-4 SaR Dec 2005 DVOR ground station Page 24 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 39 RING ANTENNAS 90 cm ∅ CENTER ANTENNA 13.5 m ∅ REFLECTOR 40m ∅ WIRE NET 39 RING ANTENNAS CENTER ANTENNA For Training Purposes Only 1.3 m 5m TRANSMITTER Figure 10 HAM US/F-4 SaR Dec 2005 DVOR Antenna installation Page 25 Part-66 Frequency spectrum of DVOR The frequency spectrum of conventional VOR and DVOR look the same, but as shown in Figure 8 and Figure 11 , the reference and variable signal are interchanged. A VOR receiver always uses the 30 Hz AM as variable phase and the 30 Hz FM as a reference phase. But the result of the bearing calculation is independent from the type of ground station, because the direction of rotation is opposite: − conventional VOR : CW − DVOR: CCW For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION VOR HAM US/F-4 SaR Dec 2005 Page 26 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 DOPPLER FREQUENCY VARIATION 480 Hz CARRIER VOICE VOICE fC fC - 1020 Hz fC + 1020 Hz IDENTIFIER fC - 9960 Hz DOPPLER FREQUENCY VARIATION 480 Hz IDENTIFIER fC - 30 Hz fC + 30 Hz fC + 9960 Hz 30 HZ AM 30 HZ FM Variable Reference Variable For Training Purposes Only 30 HZ FM Figure 11 HAM US/F-4 SaR Dec 2005 DVOR Spectrum Page 27 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 VOR RANGE AND ACCURACY Range The maximum range of the VOR is limited S due to the used VHF frequency and the altitude The maximum range is calculated by a formular given in ICAO ANNEX 10: D 1,23 h D: Distance in Nm h: Altitude of the A/C in ft S due to the Class of the VOR groundstation Range Nm For Training Purposes Only VOR Class within Altitude ft Terminal ( TVOR ) 25 1000 - 12.000 Low Altitude ( L ) 40 1000 - 18.000 High Altitude ( H ) 40 100 130 Accuracy The ICAO requirements of overall accuracy for the VOR system are 5 . The error sources can be : S ground equipment S propagation conditions S aircraft equipment S readings For the ground equipment, the errors can be produced by the antennasystem or unwanted modulation ( e.g. 30 Hz AM on subcarrier ). After Course Alignment, these errors have to be less than 3,5 in all directions. The course alignment is done every 4 month by testflights. Errors produced by bad propagation conditions like reflectors ( e.g. buildings, hills or mountains ) can be reduced by using a Doppler VOR groundstation. Under the same groundconditions, the error of a DVOR will be about 0,5 , if a conventional VOR has an error of 3,5 ( DVOR error 1/10 VOR error ). 1000 - 14.500 14.500 - 60.000 18.000 - 45.000 The radiated rf power has only small influence to the maximum range, but the antenna characteristics of the ground - and also the aircraft - antenna as well as the altitude influence the max. range. As a result of the radiation characteristic of the ground antenna - it must be horizontal polarized - there is an area overhead the station, where the received signals are not reliable. This area is called ”cone of confusion” or ”cone of silence”. In this area no VOR navigation is possible. HAM US/F-4 SaR Dec 2005 Page 28 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 Mountain reflected signal direct signal For Training Purposes Only wrong course indication due to interference of direct - and reflected - signal Figure 12 HAM US/F-4 SaR Dec 2005 VOR- Range / - Accuracy Page 29 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 VOR AIRCRAFT EQUIPMENT Aircraft Components The aircraft is normally equipped with two independent VOR systems. Each system consist of S VOR Antenna S VOR Receiver S VHF NAV Control panel S VOR Indicator For Training Purposes Only Basic function of Aircraft VOR Equipment The function of the aircraft equipment is to select and receive the signal from the wanted VOR groundstation, to process the signal and to indicate the result to the pilots and to feed it to the linked systems. The VOR process can be divided in two groups: S Automatic VOR S Manual VOR The Automatic VOR continuously indicate the VOR bearing of the selected groundstation to the pilots. The indication is done in the VOR/RMI or RDDMI or as an additional feature in the ND/EHSI. The pilots only have to select the frequency of the groundstation. The Manual VOR supplies the pilots with information about the deviation to a preselected course ( PSC / CRS ). The VOR Deviation is indicated in the HSI/EHSI/ND and also send to the FMGEC for VOR modes and for radio position calculation of the FMC part. The manual VOR produces also information, if the aircraft will fly TO or FROM the VOR station if the preselected course ( CRS ) is flown. HAM US/F-4 SaR Dec 2005 Page 30 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 VOR ANTENNA AUTOMATIC VOR/RMI VOR RDDMI Frequ. HSI / EHSI ND MANUAL VOR For Training Purposes Only CRS VHF NAV Control Panel FMGEC (Flight Management Guidance Envelope Computer) Figure 13 HAM US/F-4 SaR Dec 2005 VOR System Interface Page 31 Lufthansa Technical Training M13.3 NAVIGATION VOR Part-66 AUTOMATIC VOR The automatic VOR always indicates the Magnetic Bearing (QDM see Figure 4 ) to the selected VOR groundstation. The bearing calculation is done by comparison between the 30 Hz variable and 30 Hz reference phase. The result of the comparison will be the Radial or QDR. To get the QDM, 180 has to be added to the value of the radial. QDM = QDR + 180 To also get the direction referenced to the aircraft lubber line ( RB ), the QDM is linked to the aircraft heading ( HDG ). VOR RB = QDM - HDG The linkage can be done within the VOR receiver or within the VOR RMI. For Training Purposes Only VOR RMI / RDDMI HAM US/F-4 SaR Dec 2005 Page 32 Part-66 For Training Purposes Only Lufthansa Technical Training M13.3 NAVIGATION VOR Figure 14 HAM US/F-4 SaR Dec 2005 VOR Automatic Indications Page 33 Lufthansa Technical Training M11.5.2 / M13.3 NAVIGATION VOR Part-66 MANUAL VOR The manual VOR provides the pilots with information referenced to a preselected course ( PSC or CRS ). If the aircraft flies left or right from this selected course, this situation is shown by a Deviation - bar or - pointer on the indicator . Additional to the deviation, the pilot gets information if the selected course will lead the aircraft TO or FROM the groundstation. Modern instruments have combined the manual VOR information and the compass information in one instrument. That gives a pictorial presentation of the aircraft position referenced to the VOR station and the selected course. The instrument is therefor known under the name „ Horizontal Situation Indicator HSI „ . In the EFIS aircraft it is also called EHSI or Navigation Display ND. The preselected course can be entered at the HSI or a control panel or on a page of the MCDU. HSI / EHSI / ND For Training Purposes Only VOR Deviation To get a deviation indication in degree, the instrument is devided into dots. Usually there are two dots to the left and two dots to the right. The VOR Full Deflection is defined by ICAO ANNEX 10 to 10 ° HAM US/F-4 SaR Dec 2005 Page 34 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.3 NAVIGATION VOR Figure 15 HAM US/F-4 SaR Dec 2005 VOR Deviation Page 35 Lufthansa Technical Training M11.5.2 / M13.3 NAVIGATION VOR Part-66 TO / FROM Indication To show the pilots, where the aircraft is referenced to the preselected course, there is a TO / FROM indication. The TO / FROM indication is indipendent from the actual aircraft heading. It indicates, if the aircraft will fly to or from the groundstation, if it flies the preselected course. For Training Purposes Only TO The TO / FROM border line is automatically generated by putting a line thrugh the VOR station 90° to the preselected course. There is an allowed sector of less than 30 ° in which the TO / FROM indication is indefinite. This sector is called „ zone of confusion „. ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ ÇÇÇÇÇÇÇ FROM 30 ° Figure 16 HAM US/F-4 SaR Dec 2005 VOR TO / FROM Page 36 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.3 NAVIGATION VOR Figure 17 HAM US/F-4 SaR Dec 2005 VOR manual Indications Page 37 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 VOR ANTENNA For reception of the VOR groundstation rf signal an antenna for the frequency range 108 to 118 Mc is necessary. The antenna has to be horizontal polarized and nondirectional. The antenna impedance shall be 50 for the wanted frequency range. The VSWR for antenna and antennacable has to be less than 6:1 for the frequency range. VOR antennas, that fulfill those requirements, are „ balanced loop „ and „ balanced slot „ types. To produce a maximum horizontal antenna gain, the installation should be as high as possible. Therefor the vertical stabilizer is a common location for VOR antenna installation. The principle of the balanced antenna is to generate a nondirectional horizontal polarized diagramm by adding two radiators with the correct phasing. For Training Purposes Only Balanced Slot Antenna The antenna consist of two „ radiating „ slots each mounted on one side of the stabilizer. Both elements are linked via an antenna phasing cable. To get the necessary phaserelationship, the length of the cable must be of odd number of /2. Balanced Loop Antenna The loop−type omnidirectional antenna is horizontally polarized, has an impedance of 50 Ohm and a maximum VSWR of 2.5 : 1 when both systems are connected. It consists of aluminum elements mounted on a fibreglass shell. The aluminum elements are bonded through a coupler. The antenna is fitted on the tip of the vertical stabilizer and is capable of receiving signals in the frequency range of 108 MHz to 118 MHz. One dual−channel VOR/LOC antenna is commonly used to supply VOR ground station rf signal to the #1 and #2 VOR/MKR receivers. Very often the antenna also supplies localizer rf to the ILS receivers through a VOR/LOC rf power dividers prior to the approach mode of operation during an ILS approach. HAM US/F-4 SaR Dec 2005 Page 38 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 VOR ANTENNA RIVETED TO STRUCTURE RADIATING SLOT FIBRE GLASS COVER FWD UP COVER PLATE PHASING CABLE ANTENNA CABLES Balanced Slot LIGHTNING DIVERTER STRIP VERTICAL STABILIZER TIP TOP VIEW For Training Purposes Only + − Cover − + ANTENNA FWD VOR/LOC ANTENNA SCREWS SIDE VIEW Figure 18 Dec 2005 VOR/LOC Antenna Installation COAX CONNECTOR (2 PLACES) Balanced Loop HAM US/F-4 SaR SCREWS VOR Antenna Page 39 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION VOR Part-66 VOR RECEIVER The transmitted ground station signal is applied to RF receiver circuits via the antenna. In the RF receiver, the signal is passed through preselector to a mixer, where it is combined with the VCO injection signal from the synthesizer to produce the IF. The VCO injection signal is a function of the VOR tuning input, which is done automatically by FMGCs or manually via the appropriate MCDU page or as NAV Back Up Tuning via the RMP/RCP. After amplification the IF signal is fed to a detector, where the modulation is recovered from the IF signal to produce the NAV output. This output is applied to the audio amplifier module, where it is applied to an audio processor. The audio processor passes audio frequencies between 350 Hz and 2,500 Hz and rejects other outputs. The output is fed to the audio output and then to the aircraft audio integrating system. In modern VOR receivers, the audio is additionally fed to a Morse code detector, to supply the ARINC 429 XMTRS with VOR IDENT for indication on the NDs. The NAV signal of the receiver is fed to a bearing computer, where it is routed in two directions, to the 30 Hz variable bandpass filter and to the 9,960 Hz band− pass filter. In the first direction, all frequencies other than 30 Hz are rejected. The 30 Hz variable is fed to the phase comparator. In the second direction, the 9,960 Hz component is passed through a filter . The 9,960 Hz is then applied to a phase locked loop FM demodulator that recovers the 30 Hz reference signal. The 30 Hz reference signal is also fed to the phase comparator. The phase comparator compares the phase relationship between the two signals and converts it into ARINC 429 data bus signal. The data is transmitted on ARINC 429 bearing output data bus for application to DDRMI and DMCs ( EFIS ). HAM US/F-4 SaR Dec 2005 Page 40 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 POWER SUPPLY 115V AC FAULT MEMORY FREQUENCY SOURCE SELECT DISCRETE MCDU-FMC / RMP-RCP TUNING PORT A ARINC 429 RCVR FMC ALTERNATE PORT B ARINC 429 RCVR TO STATUS LEDS TUNING PROCESSOR TEST FAULT DATA SDI AUDIO OUTPUTS FREQUENCY SYNTHESIZER AUDIO PROCESSOR For Training Purposes Only TO ILS MORSE DECODER VOR OUTPUT 1 ARINC 429 XMTRS RECEIVER CIRCUITS RF PWR DIVIDER BEARING COMPUTER AUDIO MANAGEMENT UNIT DMC # 1/2/3 OUTPUT 2 ANTENNA VOR /MKR RECEIVER Figure 19 HAM US/F-4 SaR Dec 2005 VOR Receiver block diagram Page 41 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 VOR Outputs Today the different VOR outputsignals are normally send in ARINC 429 format. The elder VOR generations ( ARINC 547 / 579 ) deliver the outputs as analog signals. VOR Bearing The analog VOR bearing can be delivered as synchro - or sine / cosine- signal. The ARINC 547 combines the VOR bearing and the HDG within the receiver and generates a VOR RELATIVE Bearing output. The ARINC 579 combines the VOR bearing and the HDG within the VOR/RMI and generates a VOR Bearing ( QDM ) output. VOR Deviation S High Level Deviation The output is 2,00 V DC at 200 for a course sector of S Low Level Deviation The output is 150 mV DC at 200 for a course sector of VOR Termination Loads If the VOR receiver correspond with the newer ARINC 579, than the output resistor of the deviation circuit is low. The deviation outputvoltage is allowed only to deviate by 0,5% if the loadresistor changes between no load and 200 For receivers of the elder ARINC 547 the load must be exactly 200 If the load of the connected instruments and other systems ( e.g. Autoflight ) has a higher value, resistors from a network have to be added in parallel to get a 200 load Similar load switching is necessary to the TO /FROM - and Warning - outputs. 10 ° . 10 ° . For Training Purposes Only VOR To / From S High Level To / From The output current is 2 mA thru two 200 load resistors in parallel. So the output voltage is 400 mV. S Low Level To / From The output current is 200 A thru two 200 load resistors in parallel. So the output voltage is 40mV. VOR Warning S High Level Warning Valid outout is 28 V Dc. S Low Level Warning Valid outout is 250 mV Dc. HAM US/F-4 SaR Dec 2005 Page 42 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 VOR DEVIATION VOR WARNING 200 250 1000 39 30 500 48 250 49 + RIGHT 1K 40 + LEFT 41 + 1000 31 500 32 1K 1K - 42 For Training Purposes Only 43 TO / FROM 100 200 VOR / LOC DEVIATION 1K 33 44 VOR / LOC FLAG + TO 0,2K INDICATOR +FROM ARINC 547 VOR receiver Figure 20 HAM US/F-4 SaR Dec 2005 VOR Termination loads Page 43 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 VOR TUNING The VOR systems of the elder generation are tuned from a dedicated VOR control panel. The newer generation e.g. ARINC 711 , is tuned normally automatically by the FMGEC and can alternately tuned manually by the MCDU or as a Back Up via the RMP / RCP. Manual Tuning The frequency selection can be done from a combined VHF - COM / NAV or dedicated VHF - NAV control panel. For Training Purposes Only The transmission of the selected frequency is done in a standard called „ 2 out of 5 „. For the the transfer of one decade, two of five wires must be grounded. The frequency transfer is done also in ARINC 429 standard. HAM US/F-4 SaR Dec 2005 Page 44 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 Automatic Tuning S Normal tuning For aircraft equipped with Flight Management System the VOR tuning is normally done by the onside FMGEC via the RMP. The selected frequency is a function of the active company route and the PPOS. In normal operation, the RMP operates as a relay which sends the frequency/course information from the FMGEC to the VOR/MKR receiver .The FMGEC tunes the VOR/MKR receiver either automatically or manually: −tuning is automatic if no VOR is selected manually −by means of one MCDU, the pilot can select a VOR by ident or frequency Which VOR station momentarily is selected can be seen on the appropriate MCDU page and on the ND. For Training Purposes Only S Tuning in Case of Failure By a second port the VOR/MKR receiver receives a second management bus directly from the offside FMGEC. The receiver selects one of the two port functions by a discrete signal (RMP NAV DISC) which is received from the RMP. MCDU HAM US/F-4 SaR Dec 2005 Page 45 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 For Training Purposes Only THIS PAGE INTENTIONALLY LEFT BLANK HAM US/F-4 SaR Dec 2005 Page 46 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION VOR Figure 21 HAM US/F-4 SaR Dec 2005 VOR tuning Page 47 Part-66 Back Up tuning In normal configuration each FMGEC controls its on-side receivers. In case of one FMGEC failure, the remaining one controls both sides receivers. In emergency configuration (failure of the FMGEC 1 and 2) the RMP can control the VOR/MKR receiver after selection of the NAV mode on the RMP. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION VOR HAM US/F-4 SaR Dec 2005 Page 48 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION VOR Figure 22 HAM US/F-4 SaR Dec 2005 VOR Back Up tuning Page 49 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 VOR DISPLAY Manual VOR Indication Automatic VOR Indication On the navigation display (ND) On the navigation display (ND) S In ROSE−VOR mode ( Fig. A ) The data below are shown: −a dagger−shaped pointer points to the selected VOR course (cyan) (item 6) −the lateral deviation bar, which represents the VOR deviation, is shown with an arrow for the TO FROM indication (cyan) (item 2). A 2 dot deviation represents a 10 VOR deviation. −in the R top corner, the VOR characteristics come into view:VOR1 (CAPT ND) or VOR2 (F/O ND), frequency, selected course (item 4), identification (cyan). With failure of the VOR system, all the data above go out of view and the VOR1 or 2 indication turns to red (item 14). If the VOR approach is selected, the VOR APP message comes into view in the center top section of the ND (item 3). −type of station (VOR1, VOR2) −shape of the associated bearing display −station identification −mode of tuning S nothing if automatically tuned by the FMGEC, S M underlined if tuned by the MCDU and S R underlined if tuned by the RMP. A single pointer on heading dial shows the bearing of the VOR1 (item 5), a double pointer that of the VOR2 (item 1). All these data are shown in white. With failure of the VOR system, all the data above go out of view and the VOR indication turns to red (item 14). For Training Purposes Only S In ROSE−NAV and ARC modes ( Fig. B ) The characteristics and location of the VOR stations which are not already included in the flight plan, can be shown when you push the VOR−D pushbutton switch on the EFIS control panel of the FCU: −cross symbol for the VOR station (item 8) −circle plus cross symbol for the VOR/DME station (item 9). S In ROSE and ARC modes If the ADF/VOR selector switch on the EFIS control panel is set to VOR, this causes: - display of the characteristics of the VOR1 and/or VOR2 stations on the L and/or R lower corner of the ND (item 7): HAM US/F-4 SaR Dec 2005 Page 50 Lufthansa Technical Training M13.4 NAVIGATION VOR Part-66 Fig. B For Training Purposes Only Fig. A Figure 23 HAM US/F-4 SaR Dec 2005 VOR ND Indication Page 51 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Part-66 Automatic VOR Indication On the DDRMI S The DDRMI indicates the VOR bearings: S a single pointer indicates the VOR1 bearing (item 11) S a double pointer indicates the VOR2 bearing (item 10). S In case of No Computed Data (NCD), no failure flag comes into view, but the related pointer stays in the 3 o’clock position. For Training Purposes Only S A failure flag comes into view, to indicate system malfunction (item 12). In this case, the related pointer is set to the 3 o’clock position (item 13). HAM US/F-4 SaR Dec 2005 Page 52 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION VOR Figure 24 HAM US/F-4 SaR Dec 2005 VOR RDDMI Indication Page 53 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ILS Part-66 ILS GENERAL The ILS ( Instrument Landing System ) is, till today, the ICAO standardized navigation aid for landing. Especially during the critical landing phase a very accurate guidance to the runway threshold is necessary. The purpose of the ILS is to give those guidance to the pilots even under bad weather conditions. The ILS supplies the pilot with a S vertical landing plane for the lateral guidance and a S horizontal landing plane for the vertical guidance. Additionally the distance to the runway is given by two marker beacons. For the different accuracy requirements, the ICAO categorizes the ground facilities into categories. S Category I operation: A precision instrument approach and landing with a decision height not lower than 60 m (200 ft) and with either a visibility not less than 800 m or a runway visual range not less than 550 m. S Category Il operation: A precision instrument approach and landing with a decision height lower than 60 m (200 ft) but not lower than 30 m (100 ft), and a runway visual range not less than 350 m. S Category III A operation: A precision instrument approach and landing with: a) a decision height lower than 30 m (100 ft), or no decision height; and Frequ. Range : Channel : Modulation : Range Tuning : : LOC 108.10 − 111.95 Mc only at odd tenth Mc GS 329,15 − 335,0 Mc paired with LOC channels LOC 50 Kc GS 150 Kc 90 Hz AM 150 Hz AM 1020 Hz identifier, (300-3000 Hz voice) Approx. 25 NM within 10 from the front course S Manual at a VHF NAV control panel, S Automatically by FMGECs or S Manually by the MCDU’s or S by RMPs in NAV BACK UP MODE ILS landing chart Figure 25 is an example of an ILS approach for the runway 25L in Frankfurt. The upper part illustrates the lateral guidance and the lower right section the vertical guidance. The distance to the touch down is given by the two marker beacons (OM , MM) and also by the DME distance to the DME FFM ( Frankfurt ). b) a runway visual range not less than 200 m. S Category III B oporation: A precision instrument approach and landing with: a) a decision height lower than 15 m (50 ft), or no decision height; and b) a runway visual range less than 200 m but not less than 50 m. S Category III C operation: A precision instrument approach and landing with no decision height and no runway visual range limitations. HAM US/F-4 SaR Dec 2005 Page 54 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Figure 25 HAM US/F-4 SaR Dec 2005 ILS Approach Chart ( FRA ) Page 55 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 ILS PRINCIPLE The principle of the ILS function - localizer as well as glideslope - can be easily illustrated by a light beam system. For each guidance plane two spotbeams are installed. One beam has the color yellow and the other one the color blue. The two lights for the localizer are aligned to the runway front course in that way, that the blue spot is on the right- side and the yellow spot is on the left- side of the center course. − If the aircraft is too far to the right, the pilot will see a blue light; this will tell him „ fly to the left „. − If the aircraft is too far to the left, the pilot will see a yellow light; this will tell him „ fly to the right „. − If the aircraft is exactly in the extension of the runway centerline, the pilot will see a green light; this will tell him „ you are on the correct approach course „. For the glideslope or glidepath the light principle is similar but the colored spots are in the vertical aligned to the glidepath angle of about 3 . The blue spot is below the glidepath and the yellow spot is above the glidepath. For Training Purposes Only To determine the correct runway centerline or the right glideslope, two light signals are necessary. Instead of different colored lights, the ILS uses different audio modulation frequencies: − 150 Hz ¢ blue − 90 Hz ¢ yellow HAM US/F-4 SaR Dec 2005 Page 56 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 BLUE For Training Purposes Only YELLOW Figure 26 HAM US/F-4 SaR Dec 2005 ILS Principle Page 57 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 ILS GROUND FACILITIES The ILS ground station is composed of: S localizer transmitter LOC − located at the departure end of the runway S glide slope transmitter GS − located on one side of the front course of the runway S marker beacons MKR − OM located approx. 3.9 NM ( 7.2 km ) − MM located approx. 3 500 ft ( 1 050 m ) For Training Purposes Only The requirements for the ILS are fixed in the ICAO ANNEX 10. HAM US/F-4 SaR Dec 2005 Page 58 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 LOCALIZER PLANE GLIDE SLOPE PLANE RUNWAY CENTER LINE RUNWAY GLIDE SLOPE TRANSMITTER MIDDLE MARKER OUTER MARKER For Training Purposes Only LOCALIZER TRANSMITTER Figure 27 HAM US/F-4 SaR Dec 2005 ILS Ground Facilities Page 59 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 LOCALIZER THEORY Principle of Localizer Instead of light beams, the localizer is transmitting radio signals (108.10-111.95Mc) in two lobes left and right along the runway center line. Each lobe is AM modulated with an audio signal. LOCALIZER ANTENNA ARRAY RUNWAY For Training Purposes Only In direction of the front course, the right lobe is predominantly modulated with 150 Hz and the lobe to the left is predominantly modulated with 90 Hz. The antenna array may for example consist out of 13 dipoles. Each dipol is feeded individually by the transmitter. The middle part of three dipoles transmits the carrier with both 90 Hz and 150 Hz modulations. The left part is modulated with +90 Hz and –150 Hz, and the right part with –90 Hz and +150 Hz. (The signs mean phase relations of +900 and –900 .) Due to the different feeding signals, the amount of modulation degree changes with respect to the runway centerline. The modulation degree of 90 Hz decreases from left to right, while the degree of 150 Hz increases and vice versa ( see Figure 29 upper part). The sum of all antennadiagrams is the desired pattern. The line of equal modulation ( for localizer it is 20 % ) defines the localizer course. The methode of measuring the localizer course is to measure the „difference in depth of modulation (ddm)“( see Figure 29 lower part). The localizer center course is defined with a ddm = 0. The full deflection of an instrument is defined with a ddm = 15.5 % or 0.155. HAM US/F-4 SaR Dec 2005 Figure 28 LOC Antenna Array ( HAM Rwy 23 ) Page 60 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 ddm LEFT ON COURSE RIGHT 15,5 % For Training Purposes Only FULL DEFLECTION Figure 29 HAM US/F-4 SaR Dec 2005 LOC DDM Page 61 Part-66 Localizer Parameter S Frequency The LOC frequency is transmitted in the range of 108.10 Mc to 111.95 Mc with a channel spacing of 50 kHz. Only the odd tenth of Mc can be used at international airports, because the TVOR station uses the even tenth Mc. The radiated signal is horizontal polarized. Additionally to the IDENT, some ILS station transmit also a voice information. The rf output power is between 10W and 30W. S Range ( see Figure 30 left section) The necessary range is given by the requirements of the ICAO standards contained in ANNEX 10. "100 left and right of the runway center line the coverage should be 25 nM at a height of 600 m above the runway threshold. Between "100 and "350 the coverage should be 17 nM. If coverage outside this sector is provided, the range should be 10 nM. Two frequency LOC. Localizer coverage may be achieved by using two composite field patterns on different carrier frequencies within the channel bandwidth. S One field pattern gives accurate course indications within the front course sector S The other pattern provides ILS indications outside the front course to meet the coverage requirements. S DDM and Displacement sensitivity ( see Figure 30 right section) The course sector A has to be adjustable from 20 to 60. The sector is determined with a ddm of 0.155 at both sides of the course center line. The course sector is adjusted to an angle, that produces a course width of 214 m at the runway-threshold or ILS Reference Datum for a ddm of 0.155. The displacement sensitivity in this sector is = 0.00145 ddm / m. ILS REFERENCE DATUM 15 m RUNWAY THRESHOLD Within the course sector B the DDM increases linearly from zero to value of 0.180, and then > 0.180. Outside the sector B the ddm must be > 0.155 except for a back beam ( if provided ). For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ILS HAM US/F-4 SaR Dec 2005 Page 62 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ILS Figure 30 HAM US/F-4 SaR Dec 2005 LOC Range / DDM Page 63 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 S LOC Accuracy Because the coverage has to be at least "350 , the radiosignal must be transmitted also left and right off the centerline. Due to builings , hangars etc. the sum of direct- and the reflected- signal can result in Bends. BENDS RUNWAY LOC ANT BUILDING The maximum allowed amplitude of deviation for a safe landing is given by the ANNEX 10 diagram ( Figure 31 ) for the different categories. For Training Purposes Only The deviation is given in microamperes 150 A ¢ 0.155 ddm. HAM US/F-4 SaR Dec 2005 Page 64 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ILS Figure 31 HAM US/F-4 SaR Dec 2005 LOC Accuracy Page 65 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION ILS Part-66 GLIDESLOPE THEORY Principle of Glideslope Similar to the LOC system, the glideslope is transmitting radio signals ( 329,15 − 335,0 Mc ) - instead of light beams - but the lobes are in the vertical plane above and below the glidepath. Above the glidepath, the lobe is predominantly modulated with 90 Hz and the lobe below the glidepath is predominantly modulated with 150 Hz. The antenna array is located about 150 m left or right of the runway at about the approach threshold. Especially for the glideslope antenna, the reflection constancy of the area in front of the antenna is very important. The array comprises two or three antennas to generate the glide path transmission pattern with 90 and 150 Hz. The antennas are feeded individually by the transmitter. S The lower antenna transmits the carrier plus the sidebands +90 and +150 Hz. S The upper antenna transmits only the sidebands –90 and +150 Hz. The sum of both antenna diagrams is the desired glidepath pattern. Due to the different feeding signals, the amount of modulation degree changes with respect to the glidepath. The modulation degree of 90 Hz decreases from above to below, while the degree of 150 Hz increases and vice versa ( see Figure 33 ). The line of equal modulation ( for glideslope it is 40 % ) defines the glidepath. The methode of measuring the glidepath is to measure the „difference in depth of modulation (ddm)“. HAM US/F-4 SaR Dec 2005 The glidepath is defined with a ddm = 0. The glideslope full deflection of an instrument is defined with a ddm = 17.5 % or 0.175. ( The diffrence to LOC results by the different modulation degrees at ddm=0. ) Figure 32 GS Antenna Array If only two antennas are used, an additional false inverted glide path is generated above the normal glide path. Page 66 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 3 5 = Glidepath For Training Purposes Only Figure 33 HAM US/F-4 SaR Dec 2005 Elevation Angle = false inverted Glidepath = false Glidepath GS DDM Page 67 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 GS Parameter S Frequency The GS frequency is transmitted in the range of 329,15 − 335,0 Mc with a channel spacing of 150 kHz. The radiated signal is horizontal polarized. There is a fixed linkage between LOC- and GS-frequency; so the frequency selection is only done via the LOC frequency and GS is tuned automatically. For Training Purposes Only Localizer S Range ( see Figure 34 left) The necessary range is given by the requirements of the ICAO standards contained in ANNEX 10. S Glidepath Angle ( see Figure 34 right) For a sufficient safety height, the glidepath angle has to be adjustable between 2 and 4. The necessary angle is achieved by different installation heights of the antenna elements. Glide Slope Localizer Glide Slope MHz MHz MHz MHz 108.10 334.70 110.1 334.40 108.15 334.55 110.15 334.25 GS Angle 108.3 334.10 110.3 335.00 2,5 5.2 m 10.40 m 108.35 333.95 110.35 334.85 3,0 4.34 m 8.68 m 108.5 329.90 110.5 329.60 3,5 3.72 m 7.44 m 108.55 329.75 110.55 329.45 108.7 330.50 110.70 330.20 108.75 330.35 110.75 330.05 108.9 329.30 110.90 330.80 108.95 329.15 110.95 330.65 109.1 331.40 111.10 331.70 109.15 331.25 111.15 331.55 109.3 332.00 111.30 332.30 109.35 331.85 111.35 332.15 109.50 332.60 111.50 332.9 109.55 332.45 111.55 332.75 109.70 333.20 111.70 333.5 109.75 333.05 111.75 333.35 109.90 333.80 111.90 331.1 109.95 333.65 111.95 330.95 HAM US/F-4 SaR Dec 2005 h1 h2 The sector width from full up to full down is defined with a ddm of 0.175. With an angle of 3, the sector width is about "0.5. Page 68 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 For Training Purposes Only Glidepath adjustable between: Figure 34 HAM US/F-4 SaR Dec 2005 GS Range / DDM Page 69 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 S GS Accuracy Because the Glide Slope antenna diagram is strongly influenced by the reflectivity characteristic of the ground, there are critical and sensitive areas around the antenna. The critical area is an area where vehicles, including aircraft, are excluded during all ILS operations. This area is about 250 m from the antenna to the runway. The sensitive area is an area extending beyond the critical area where parking and/or movement of vehicles is controlled during ILS operations. It extends to about 1 km in front of the antenna. The maximum allowed bends for a safe CAT I / II / III landing is given by ICAO ANNEX 10 diagram ( Figure 35 ). For Training Purposes Only The deviation is given in microamperes 150 A ¢ 0.175 ddm. HAM US/F-4 SaR Dec 2005 Page 70 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ILS Figure 35 HAM US/F-4 SaR Dec 2005 GS Accuracy Page 71 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 ILS AIRCRAFT EQUIPMENT Aircraft Components The aircraft is normally equipped with two or sometimes three independent ILS. Each system consist of S LOC Antenna S GS Antenna S ILS Receiver ( MMR ) S VHF NAV Control panel S ILS Displays For Training Purposes Only Basic function of Aircraft ILS Equipment The function of the aircraft equipment is to select and receive the signal from the wanted ILS groundstation, to process the signal and to indicate the result to the pilots and to feed it to the linked systems. If the aircraft is equipped with an autopilot system ( FMGEC ), the ILS signals can be used for an automatic landing. The ILS signal is used for the EGPWS for Mode 5 and the envelope modulation function. The ILS process can be divided in two groups: S Localizer S Glideslope The indication of the ILS signals is done on the PFD as well as on the ND. The minimum requirements for LOC are described in RTCA DO 131 and for GS in DO 132. HAM US/F-4 SaR Dec 2005 Page 72 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 PFD For Training Purposes Only ND ( MMR ) FMGEC EGPWS Figure 36 HAM US/F-4 SaR Dec 2005 ILS System Page 73 Part-66 ILS ANTENNAS LOC Antenna The requirements for LOC- and VOR - Antenna are more or less the same, so that very often the LOC system uses the VOR antenna. The polarization must be horizontal, but there is no request for a non directional diagramm ( as it is for VOR ). Modern large airplanes are using very often a seperate LOC antenna, located under the radom. G/S Antenna For glide slope reception an antenna for a higher frequency range of 329,15 – 335,0 Mc is required. The antenna input- or feeding- resistor shall be 52. For the total frequency range the VSWR shall be less than 5. The polarization is horizontal and there is no request for a non directional diagramm. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS HAM US/F-4 SaR Dec 2005 Page 74 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 RIGHT GLIDE SLOPE TRACK ANTENNA ACCESS PANEL RIGHT GLIDE SLOPE TRACK ANTENNA SEE SEE A A FWD COAXIAL CONNECTOR RIGHT AFT NOSE GEAR DOOR WEATHER RADAR ANTENNA (EXAMPLE) DUAL LOCALIZER ANTENNA For Training Purposes Only DUAL GLIDE SLOPE RIGHT GLIDE SLOPE TRACK ANTENNA CAPTURE ANTENNA RIGHT GLIDE SLOPE TRACK ANTENNA (REMOVED) A Localizer / GS Capture Antenna Installation Glide Slope Track Antenna Installation Figure 37 HAM US/F-4 SaR Dec 2005 ILS Antennas Page 75 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 The deviation is send in BNR format by bits 17 to 28. The bit 28 represents the ddm 0,2. The maximum value of all bits is 0,4. ILS RECEIVER The ILS receiver in modern installations contains the LOC and G/S reception parts as well as the processing circuits for LOC and G/S. The ILS can also be part of a Multi Mode Receiver ( MMR ) which contains all possible landing systems like ILS, MLS and GLS (GPS ). The DDM is used to drive the deviation bars, and the sum of the 90 / 150 Hz modulation signals is used for the valid- or flag- indication. Localizer Warning The LOC signal must be INVALID under the following conditions: S No RF signal or no 90 Hz and 150 Hz modulation. S Missing 90 Hz or 150 Hz modulation and the degree of modulation of the remaining is only 20%. S The rf signal of a standard deviation signal ( 0,093 ddm ) results only in 50% deflection of the standard deviation ( 60% ). The Flag- or Valid- signal can be delivered as High-, Low- Level or ARINC 429 warning signal. S High Level output: 28 Volt DC Valid S Low Level output: 250 mV DC on termination resistor of 250 S ARINC 429 output: Sign Status Matrix SSM For Training Purposes Only Glide Slope Deviation The pointer deviation is determined by ICAO ANNEX 10. The norm for FULL Deviation is: 0,175 ddm. The norm for a STANDARD G/S Deviation is determined at a ddm of 0,092. The deviation outputsignal can be delivered as High-, Low- Level or ARINC 429 signal. Localizer Deviation The pointer deviation is determined by ICAO ANNEX 10. The norm for FULL Deviation is: 0,155 ddm. The norm for a STANDARD LOC Deviation is determined at a ddm of 0,093. This standard deviation results in a deflection of 60% or 3/5 of full deviation. The deviation outputsignal can be delivered as High-, Low- Level or ARINC 429 signal. S LOC High Level output ( ARINC 578 ) "2 Volt on termination resistors between 200 and R for ddm= 0,155. S LOC Low Level output ( ARINC 547 ) "150 mV on termination between 200 and R for ddm= 0,155. S ARINC 429 output HAM US/F-4 SaR Dec 2005 Glide Slope Warning The G/S signal must be INVALID under the following conditions: S No RF signal or no 90 Hz and 150 Hz modulation. S Missing 90 Hz or 150 Hz modulation and the degree of modulation of the remaining is only 40%. S The rf signal of a standard deviation signal ( 0,092 ddm ) results only in 50% deflection of the standard deviation. The Flag- or Valid- signal can be delivered as High-, Low- Level or ARINC 429 warning signal. LOC – G/S Termination Resistors Units of the elder standard ARINC 547 need a defined termination resistor. If the resistor of the connected instruments and/or units differ from the required one, dummy resistors must be used. Page 76 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION ILS Part-66 Deviation from Average in Modulation degree Indication Note Note: X definition according to ICAO ANNEX 10 Figure 38 HAM US/F-4 SaR Dec 2005 ILS Standard Deflection Page 77 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 ILS TUNING The ILS systems of the elder generation are tuned from a dedicated NAV control panel. The newer generation is tuned normally automatically by the FMGEC and can alternately tuned manually by the MCDU or as a Back Up via the RMP / RCP. Manual Tuning The frequency selection can be done from a dedicated NAV control panel. For Training Purposes Only The transmission of the selected frequency is done in a standard called „ 2 out of 5 „. For the the transfer of one decade, two of five wires must be grounded. The frequency transfer is done also in ARINC 429 standard. The Glide Slope frequency is automatically tuned with the Localizer. HAM US/F-4 SaR Dec 2005 Page 78 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 Automatic Tuning S Normal tuning For aircraft equipped with Flight Management System the ILS tuning is normally done by the onside FMGEC via the RMP. The selected frequency is a function of the active company route, SIDS, STAR and the PPOS. In normal operation, the RMP operates as a relay which sends the frequency/ course information from the FMGEC to the ILS receiver . The FMGEC tunes the ILS receiver either automatically or manually: −tuning is automatic if no ILS is selected manually −by means of one MCDU, the pilot can select a ILS by ident or frequency. Which ILS station momentarily is selected can be seen on the appropriate MCDU page, on the PFD and on the ND. USE TO HFW 108 05 ENTER ILS DATA 232 S Tuning in Case of Failure By a second port the ILS receiver receives a second management bus directly from the offside FMGEC. The receiver selects one of the two port functions by a discrete signal (RMP NAV DISC) which is received from the RMP. For Training Purposes Only Back Up tuning In normal configuration each FMGEC controls its on-side receivers. In case of one FMGEC failure, the remaining one controls both sides receivers. In emergency configuration (failure of the FMGEC 1 and 2) each RMP can control both ILS receiver after selection of the NAV mode on the RMP. MCDU HAM US/F-4 SaR Dec 2005 Page 79 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 For Training Purposes Only THIS PAGE INTENTIONALLY LEFT BLANK HAM US/F-4 SaR Dec 2005 Page 80 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ILS Figure 39 HAM US/F-4 SaR Dec 2005 ILS tuning Page 81 Lufthansa Technical Training M13.4 NAVIGATION ILS Part-66 ILS DISPLAYS Localizer The LOC is indicated on the PFD ( ADI / EADI ) and / or ND ( HSI / EHSI ). The normal indication of LOC #1 is on PFD#1 and ND#2 and for LOC#2 vice versa. The deflection of the bar indicates the deviation from the runway centerline course. The bar moves to the left, if the A/C is within the 150 Hz ( blue ) beam and moves to the right if the A/C is within the 90 Hz ( yellow ) beam. The deviation at ILS is independent from the selected runway heading. A Fulldeflection of the LOC-deviation bar from center to right or left is given at a ddm of 0,155. As a function of the runway length, this represents an angle of "1 to "3 In some installations, the LOC scale in the PFD can be an Expanded Scale. For Training Purposes Only Back Beam Back Beam approaches are no more common or standard, but can be used. In this case the aircraft is approaching from the opposite direction of the normal ILS beam. The pilots have to take into consideration, that now 150 Hz is left and 90 Hz is on the right and that the deflection of the indication will go to the ”wrong” direction. It shows the position of the aircraft related to the beam. To overcome this confusion (accidents had happened), the pilot selects the front course for the ILS while approaching on the back beam. This trick makes the HSI to a correct situation indicator again and the deflection is now to the correct side. The glideslope indication is of no use on a back beam approach and has to be neglected. PFD LOC scale ILS: - Ident - Frequ. SEL RWY HDG ND ddm 0,155 SEL CRS RWY HDG LOC DEVIATION BAR HAM US/F-4 SaR Dec 2005 Page 82 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ILS Figure 40 HAM US/F-4 SaR Dec 2005 LOC Indication front beam Page 83 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Part-66 Gide Slope The G/S is indicated on the PFD ( ADI / EADI ) and / or ND ( HSI / EHSI ). The normal indication of G/S #1 is on PFD #1 and ND #2 and for G/S #2 vice versa. The deflection of the bar indicates the deviation from the glidepath centerline. The bar moves up, if the A/C is within the 150 Hz ( blue ) beam and moves down, if the A/C is within the 90 Hz ( yellow ) beam. A Fulldeflection of the G/S-deviation bar from center to up or down is given at a ddm of 0,175. As a function of the glidepath angle, this represents an angle of "0,5 to "1 PFD GS scale ILS: - Ident - Frequ. ND For Training Purposes Only excessive ddm ddm 0,175 HAM US/F-4 SaR Dec 2005 Page 84 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ILS Figure 41 HAM US/F-4 SaR Dec 2005 G/S Indication Page 85 Part-66 MARKER SYSTEM The marker system is a radio navigation aid. It is normally used together with the ILS during an ILS approach. The markers are also used on airways, to mark a position on airways. In Germany there are no Airway Marker any more. Two types of markers are used: S Z-Marker S Fan-Marker The Z- Marker whose vertical beam is cone shaped, is to mark a certain crossing point of airways, or to mark the cone of silence over a NDB. The Fan- Marker transmits a fan shaped - elliptic or boneshaped- vertical beam to mark the position along the airways or to give a distance to the runway. The system provides visual and aural indications. All marker transmitters transmit a modulated 75 MHz signal to provide a marker position. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION MARKER HAM US/F-4 SaR Dec 2005 Page 86 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION MARKER Part-66 FAN-MARKER bone shaped For Training Purposes Only FAN-MARKER elliptic Z-MARKER circular Figure 42 HAM US/F-4 SaR Dec 2005 Marker System Schematic Page 87 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION MARKER Part-66 There can be three ILS marker transmitters positioned on the ground at known distances from the runway threshold: S the outer marker at approx. 4 NM S the middle marker at approx 0.5 NM S the inner marker- if available- at approx 0.1 NM from the runway threshold. In some countries airway markers are using the inner marker mod. frequency. When the aircraft passes through the beam of a marker transmitter, the modulating frequency is detected and the system provides aural and visual indications to the flight crew. Indication Outer Marker : “ OM ” − BLUE − 400 Hz−Modulation - dashes ----Middle Marker : “ MM ” − AMBER− 1300 Hz−Modulation - dots and dashes .-.Inner Marker : “ IM ” − WHITE− 3000 Hz−Modulation - dots ..... For Training Purposes Only Receiver The marker receiver may be an extra black box, or - in modern aircrafts - a combined VOR / Marker receiver. HAM US/F-4 SaR Dec 2005 Page 88 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION MARKER Figure 43 HAM US/F-4 SaR Dec 2005 Marker Ground Station Page 89 Part-66 MKR Antenna Marker beacon transmit horizontally polarized signals on a frequency of 75 Mc. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION MARKER Figure 44 HAM US/F-4 SaR Dec 2005 MKR Antenna Page 90 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION MARKER Part-66 Marker Indication The Marker indication can be done by colored lights for each type of marker or by just one light labeled MKR. Some elder aircraft installation might have a Marker HIGH / LOW switch to increase the normal receiver sensitivity for detection of airway marker. For ILS approaches the switch must be in LOW position to get the exact passing position of the OM or MM. For Training Purposes Only In „ glass cockpits „ the marker indication may be integrated in the EADI or PFD. Figure 45 HAM US/F-4 SaR Dec 2005 Marker Indication Page 91 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION MLS Part-66 MLS GENERAL The MLS ( Microwave Landing System ) was scheduled to be implemented progressively until about 1999. It is intended supplant the Instrument Landing System (ILS) as the international standard landing aid. Till now ( 2001 ) MLS has been demonstrated to be superior to ILS from both performance and reliability stand points. Yet considerable debate persists on its relative merits. Although tolerance criterion for flight safety are imbedded in these comparisons, flight safety has not been a major component of these discussions. The effect of MLS on airport capacity is a central issue. The Microwave Landing System has become the most controversial aspect of the FAA’s National Airspace System (NAS) plan. Since inception of the MLS program, nearly 25 years ago, the FAA has taken the lead in developing and evaluating MLS. The FAA has undertaken a broad range of projects to insure MLS will be available to meet precision landing needs of the future. After two decades of effort, consensus on the value of MLS has failed to materialize. In fact, support has declined in recent years, as efforts to develop and test operational units have undergone several major set,backs. The benefits of MLS are generally related in terms of the availability of : S curved approaches, S wide area coverage, S high signal integrity, S back azimuth guidance as well as S high reliability and ease of maintenance. Unfortunately, the most advantageous elements are the most difficult to implement. The curved approach, which is the most widely advertised and useful benefit, is the furthest from implementation. Curved approaches require the addition of a navigation computer and Terminal Instruction Procedures (TERPS) that are still under development. Based on the difficulty that Hazeltine Corporation has had in developing Category I units (no curved approach), the aviation community is uneasy about the technical complexity and cost of realizing MLS’s full benefits. As a result, efforts to generate commitment to MLS have faltered. HAM US/F-4 SaR Dec 2005 Page 92 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION MLS ILS MLS Figure 46 HAM US/F-4 SaR Dec 2005 MLS and ILS Page 93 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION MLS Part-66 Historical Backgrounds ILS The first ”precision approach” landing aid was the instrument landing system (ILS). In 1939, the Civil Aeronautics Authority (CAA) demonstrated the first commercial ILS in Indianapolis, Indiana. In 1941, the system was adopted as the national standard, and in 1949, the ICAO adopted it as the international standard. ILS has undergone several enhancements but remains essentially unchanged in its basic principles. Under Instrument Flight Rule (IFR) landing conditions, aircraft are guided by ILS at most major airports. This involves flying down a narrow beam transmitted at a constant azimuth. Aircraft fly down with a specified separation between 4 and 6 NM to the appropriate decision height (Category I − 200 ft, Cat II − 100 ft, Cat III − 0 ft ). Although there have been persistent concerns about its signal quality, ILS has met the precision landing needs of aviation for nearly 50 years. Over the years, enhancements have achieved lower operational minimums. These enhancements have served to reduce the urgency of developing a new landing aid, but are unlikely to preclude it. Several factors prevent ILS from meeting the needs of aviation indefinitely. S Electromagnetic interference of the ILS signal is increasingly a problem, particularly from powerful FM stations but also from TV stations. S Weather has a detrimental impact on the ILS signal. S Taxiing aircraft must be routed to avoid interference with the ILS signal. S In addition, ILS must be located on flat terrain, since the surrounding area contributes to the formation of the beam. Importantly, ILS is usually the limiting item to airport capacity during IFR operations. Although an ILS malfunction has never lead to an accident, secondary contributions due to lack of availability and capacity need to be considered. The enhancements of ILS have not met the need for increasing capacity. Even with enhanced ILS, we can see empirically that delays are worsening. HAM US/F-4 SaR Dec 2005 MLS Work on MLS can be traced to shortly after the creation of the FAA in 1959. At that time, a review of existing efforts revealed that the microwave scanning beam technique appeared to have the most promise. This initial effort led to the development and delivery in 1965 of an advanced integrated landing system (AILS−microwave scanning beam system). Where ILS is a pencil beam at constant azimuth, MLS has two scanning beams one in azimuth and one in elevation. MLS has the potential to provide precision guidance to a large volume of airspace. The current MLS approach is an “air derived” system. The receiver/processor receives each signal independently and calculates the angles of elevation and azimuth. The angles are based on the difference between the arrival of signals from the “to” and “fro” scans of the beam. These angles, along with distance to go information from the Precision Distance Measuring Equipment (PDME) and auxiliary data (runway condition, minimum glide path, facility identification, operational status), are used to guide initial and final approaches. The publicized curved approaches requires additional processing of data by a navigation computer (FMS, FMGCS, etc.). The MLS has an operating frequency in the microwave C−band (5 GHz). This frequency range allows the system to utilize a smaller antenna, achieve highly directed scanning beams, and virtually eliminate the impact of weather degradation on the signal. The higher frequency also permits a higher sampling rate, and thus greater precision. Accuracies achieved by MLS can be on the order of a few tenths of a degree. Actual error, from the flight path, is a function of distance to go. Lateral and longitudinal touchdown error is a function of wind and flight control errors, in addition to MLS guidance errors. Page 94 Lufthansa Technical Training M13.4 NAVIGATION MLS Part-66 For Training Purposes Only Azimuth Elevation Figure 47 HAM US/F-4 SaR Dec 2005 MLS - ILS Azimuth / Elevation Page 95 Lufthansa Technical Training M13.4 NAVIGATION MLS Part-66 GROUND STATIONS There are basic components installed at an airfield offering basic MLS. They are: S azimuth station, S precision distance measuring equipment (PDME), S elevation station, S basic data words transmission facility. The basic configuration may be expanded to include one or both of the following: S back azimuth equipment, S flare elevation equipment. − Azimuth station − Back azimuth station The back azimuth station is identical to the azimuth station and is used to provide guidance for departures and missed approaches. The ICAO specified minimum range for the back azimuth is 7 NM, however the azimuth and back azimuth stations can interchange functions when the opposite approach direction is in use. Hence to provide this bidirectional facility both stations must have 20 nm range. Two elevation stations are needed for bidirectional capability. Only one would be switched on at a time. For Training Purposes Only The azimuth station provides a service such as that provided by an ILS localiser but with much wider cover, being + 40 degrees from the centre line extending to a range of 20 nm. It is usually positioned 1000 feet beyond the rollout end of the runway but various options do exist if conditions do not permit the normal siting. The azimuth accuracy requirement at the MLS reference datum, located 50 feet above the runway threshold, is + 20 feet. − PDME The PDME is co−located with the azimuth station and provides continuous, omnidirectional and accurate range information to an aircraft during its approach and landing phase. The accuracy available is one of two standards laid down by ICAO. Standard 1 is 100 feet and Standard 2 is 40 feet, compared with + 600 feet for the current standard DME. The transponder is compatible with the current standard DME (DME/N) but, of course, the laid down high accuracy standards are only available if a DME/P interrogator is fitted to the aircraft. Use of DME means that ILS markers are not required. − Elevation station The elevation station is comparable to an ILS glideslope facility and is located to one side of the runway, offset from the centreline by at least 250 but not more than 500 feet. The distance from the runway threshold is between 400 and 1000 feet depending on the minimum glideslope angle provided. Guidance on glidepath angles up to 15 degrees is available. The elevation accuracy requirement at the MLS reference datum is + 2 feet. Flare elevation covers the last five miles of an approach with better coverage nearer the horizontal. HAM US/F-4 SaR Dec 2005 Page 96 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION MLS Figure 48 HAM US/F-4 SaR Dec 2005 MLS Ground Stations Page 97 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION MLS Part-66 MLS PRINCIPLE The MLS operates on 200 channels in the frequency band 5031 MHz to 5091 MHz, hence we are dealing with a C−band system. All components transmit on the single frequency allocated to an approach in a time multiplexed signal format. In addition to the elevation and azimuth transmissions a burst of data is transmitted. The time reference scanning beam (TRSB) signal format is characterized by the time division multiplexing (TDM) of information. Each guidance function (azimuth, elevation, flare, back az, and aux. data) is given an independent time slot. Ground equipment transmits signals that scan vertically for elevation and horizontally for azimuth. In addition, a preamble for each time slot is transmitted via a broad beam signal throughout the coverage volume. Separate channels (200) allow airborne equipment to differentiate between runways and airports in the same local airspace. S TO - FROM Scan Only after the above preamble and sector signals have been transmitted are the pulses radiated which are used to measure the azimuth (or elevation) approach angle. There are two pulses involved, a TO and a FRO. In fact what we have is a narrow beam which is electronically swept, first in the TO direction and then in the FRO direction. Any aircraft within the coverage of the scanning beam will thus receive two pulses of energy in one complete cycle of the scan. If the two received pulses are well spaced then the aircraft will be in the left hand half sector of the cover, on the other hand if the TO and FRO pulses are close together then the aircraft will be in the right hand half sector . Note that even if the aircraft is on the extreme right of the sector of cover, the TO and FRO signals will still be separated. Transmitting Format S Preamble The preamble, which is transmitted from a wide beam aerial at the beginning of each block, allows time for RF carrier acquisition and also contains coded information relating to timing and identification of the block. The first part of this preamble is a burst of unmodulated carrier which allows for phase synchronisation, necessary since phase shift keying is used for the encoded digital data subsequently transmitted. After the unmodulated carrier a code is transmitted which allows a time reference to be established in the airborne receiver. The final part of the preamble is a function code which identifies the transmission block as being, in this case, azimuth. S Sector signal The sector signal also has three distinct parts to it. First we have part of a four character identification code for the particular station; this serves the same function as an ILS ident. The complete identification code is built up over several blocks of transmission. After the ident code a constant amplitude limited duration signal is transmitted which can be used by the aircraft installation to decide which of its aerials ( antennas ) is receiving the strongest signal. Both of these first two sector signals are transmitted from a wide beam aerial. Finally we have the azimuth out of coverage indicator pulses which are transmitted from different aerials in all directions not covered by the azimuth (or elevation) beam. HAM US/F-4 SaR Dec 2005 Page 98 Lufthansa Technical Training M13.4 NAVIGATION MLS Part-66 ÇÇ ÉÉÉ TIME ( sec) ÇÇ ÉÉÉ ËËËËËËËË ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÇÇ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉ ËËËËËËËË ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÇÇ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉ ËËËËËËËË ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÇÇ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉ ËËËËËËËË ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÇÇ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉ „ „ ËËËËËËËË ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÇÇ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉ ËËËËËËËË ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÇÇ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉ „TO“ Pulse Preamble For Training Purposes Only Ref Time „FRO“ Pulse Sector Signals TO“ SCAN TIME SLOT FRO“ SCAN TIME SLOT 0 Pause Time 0 T0 Tm Midscan Point Figure 49 HAM US/F-4 SaR Dec 2005 MLS Measurement Principle Page 99 Part-66 Time Multiplexed Transmissions Although the explanation of the basic principle of angle measurement has been in terms of azimuth, elevation or flare works in a similar way. As the transmission is done on a common frequency for azimuth-, elevation-, flareangle measurement and auxiliary data, a timetable with fixed intervals for each transmission block is necessary. Figure 50 shows a frame period of 84 msec for the time multiplexed transmissions. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION MLS HAM US/F-4 SaR Dec 2005 Page 100 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION MLS Figure 50 HAM US/F-4 SaR Dec 2005 MLS Transmission Format Page 101 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION MLS Part-66 AIRCRAFT EQUIPMENT The operational capabilities of MLS depend on the type of equipment installed in the aircraft. The basic fit allows for flying along a DME arc followed by a straight−in approach at a selected glidepath angle. This is illustrated in the Figure 51 A. The equipment needed in this case being: S MLS receiver S DME/P S control panel S HSI S range display The next level of sophistication would allow for a segmented approach under the guidance of the MLS as shown in the Figure 51 B. The equipment needed in this case being: S MLS receiver, S DME/P S control panel, S computer, S flight director, S display unit. Finally, we utilise the full capability of the MLS by being able to execute curved approaches under the guidance of the MLS shown in the Figure 51 C. The equipment needed in this case being: S MLS receiver, S DME/P S control panel, S flight management system S flight director or EFIS All three illustrations show the need for the three basic items of an MLS, namely the MLS receiver, a DME/P interrogator, and a control panel. The MLS receiver will have to process the signals from the ground MLS transmitters and calculate steering signals which can be fed to, say, a HSI in order to direct the pilot to fly left/right, up/down and so attain his selected approach path. This capability, within the receiver itself, is limited to a straight line approach. HAM US/F-4 SaR Dec 2005 The control panel will allow the pilot to select the frequency of the facility he is approaching, the required azimuth course and the elevation angle. The DME/P will be tuned by the combined controller in the same way as a DME/N is tuned by an ILS/VOR controller. The addition of a computer external to the computing facilities of the MLS receiver will allow more sophisticated approaches. How sophisticated depends on the power of the computer. The addition of a facility to make segmented approaches requires more sophistication of the controls. Waypoints are set up which define how the aircraft will approach the runway, the extra controls allow the input of these waypoints. The addition of such a controller, often as part of a combined display/ controller, is not new since the entry of waypoints is part and parcel of area navigation facilities. What is new is the use of waypoints in conjunction with a landing aid. More computing power allows curved approaches through the selected waypoints to be flown. In such cases the computer must define the turn path with a specified radius of turn. The last figure reflects the fact that where we have such sophistication then invariably we will have a powerful flight management system (FMS) which has many inputs and will provide automatic control over the aircraft’s flight controls and engine settings. We also need to match the computing power with sophisticated display systems such as electronic flight instrumentation (EFIS). Any of the configurations illustrated can include an autopilot to provide the additional capability of making automatic landings. The category of autoland will depend on the degree of redundancy in both the ground and air facilities. An ILS based system could not match the capabilities of any of the illustrated systems. Dealing with the least sophisticated the pilot is able to select the required azimuth and elevation approach angles within the limits of the coverage of the ground facility. With an ILS based system these angles are fixed. Also the region of cover is significantly greater with MLS. Page 102 Lufthansa Technical Training M13.4 NAVIGATION MLS Part-66 C A HSI For Training Purposes Only B Figure 51 HAM US/F-4 SaR Dec 2005 MLS Approach Capabilities Page 103 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ADF Part-66 AUTOMATIC DIRECTION FINDER GENERAL The automatic direction finding (ADF) system is a radio aided navigation that receives radio signals from ground stations - Non Directional Beacons ( NDBs ). It provides bearing information to an indicator and an aural output to the flight interphone system. Even if the ADF is a very old radio nav system, it is still used for S en route navigation S NDB approaches and S sometimes holdings. The specification for ADF operation is given in ICAO Annex 10 Volume I, Chapter 3.4. The frequency is within LF- and MF- band ( 190 KHz − 1750 KHz ). The band includes also the standard AM radio broadcast band as well as the low frequency radio range band from which pilots can listen to broadcasts of weather reports. The ADF aircraft system receives an electromagnetic wave transmitted by a ground station. A loop antenna intercepts the magnetic portion and couples this signal through the quadrantal error corrector to the ADF receiver. A sense antenna intercepts the electric portion and couples this signal through the sense antenna coupler to the ADF receiver. The receiver processes both signals and, if in the ADF mode, produces a signal to position the bearing pointer. The receiver also produces the audio signal to the interphone system. Ground stations are called either S Non Directional Beacons (NDB) for long range enroute navigation, transmitting with high power, or S Locators, if they are used for approaches. Locators are situated close to the marker beacons. The frequency range of NDB’s and Locators is from 200 kHz to 526.5 kHz. The extended frequency range of the ADF receiver allows to use also normal MW radio stations for navigation, if their position is known. HAM US/F-4 SaR Dec 2005 The principle of the ADF navigation is to determine the relative bearing of a selected ground station. This is obtained by the combination of: − the signals from two loop antennas positioned 90 deg. apart − the signal from an omni−directional sense antenna. This signal is not affected by the relative bearing. An additional Morse signal is provided to identify the selected ground station. The modulation of the NDBs can be A0 / A1 or A0 / A2. For A0 / A1 modulation, a BFO is necessary to listen to the IDENT. Frequency Range : Range Tuning : : 190 - 1750 KHz (Broad Casting AM - Stations) 40 - 80 Nautical Miles S manual at the ADF control panel / RMP/ MCDU or S automatically by the FMC Page 104 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Part-66 For Training Purposes Only ADF in A/C NDB, LOCATOR or MW RADIO STATION on GND Figure 52 HAM US/F-4 SaR Dec 2005 ADF - NDB Page 105 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ADF Part-66 GROUND STATIONS Three types of Ground stations can be used for ADF navigation: S Locators for approaches ( located near markers ) S Nondirectional Beacons for long range navigation S Public Radio Stations if position is known In the Figure 53 are two locators ( FW 382 and FR 297 ), collocated with the outer marker beacons. The two letters for each station are the morse code identifier and the numbers are the frequency in kHz. Figure 53 Frequency information are in the approach charts in the enroute maps. HAM US/F-4 SaR Dec 2005 In the map Figure 54 dotted circles. Locators on Approach Maps non directional beacons (NDBs) are marked as concentric Page 106 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Figure 54 HAM US/F-4 SaR Dec 2005 NDBs in Jeppesen Maps Page 107 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 ADF PRINCIPLE The ADF is a form of ‘radio compass’ that provides the pilot with the relative bearing of the beacon to which the equipment is tuned. Almost everybody knows that the reception intensity of a home radio can be varied by turning arroud the Loop Antenna (or the radio with the built in antenna). This effect is used by the ADF. A loop antenna receives the magnetic part of the electromagnetic wave as : S a maximum, ADF Direction Determining When turning arround the loop antenna, one can find out, that the increase and decrease of the signal strength arround the maximums are more widespread than arround the minimums. That means, that a bearing information based on minimum selection is more precise, than based on maximum selection. So the ADF uses the minimums. if the axis of the coil is perpendicular to the direction to the station or S a minimum For Training Purposes Only if the axis of the coil is into the direction to the station HAM US/F-4 SaR Dec 2005 Page 108 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 2 1 1 2 2 1 NDB LOOP ANTENNA MAXIMUM SIGNAL NDB MINIMUM SIGNAL 1 2 Umin Umax Umin Umax NON DIRECTIONAL For Training Purposes Only BEACON Figure 55 HAM US/F-4 SaR Dec 2005 ADF Loop Antenna Function Page 109 Part-66 Bearing Ambiguity The ear can distinguish different loudness and different frequencies, but it is no sensor for phases. Therefore if we use our home radio example again, we can find out, if the radio is slowly turned around by 3600, there will be two best positions and two worst positions for reception. Between a best and a following worst position is aN Angle of 900. Between the two best as well as between the two worst positions is an angle of 1800. Our ear is not able to find out any difference between the two opposite directions. In the same way, an ADF receiver has also an ambiguity of 1800, because with the loop antenna alone, it is not able to sense the phase. It needs a reference phase. This reference phase is delivered by a so called sense antenna, which receives the electro part of the electromagnetic wave. Overcome of the Bearing Ambiguity ( see Figure 56 ) An electromagnetic wave has a travelling direction defined by the direction of the Pointing vector. Perpendicular to that direction and to each other are the E-field and the H-field (in vaccuum or air), the electrical and the magnetic part. The H-field is sensed by the loop antenna, and the E-field is received by the sense antenna. The E-field is vertically polarized, and the received signal will not change phase, in which direction ever the aircraft will fly in azimuth. The H-field has an opposite phase between aircraft A and aircraft B. That is an unambigous criterion for the aircrafts ADF, that it is flying to or from the station. The result is, that the ADF needel in the RMI is always pointing to the station. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 110 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 PO I N T I N G E C T O R For Training Purposes Only V Figure 56 HAM US/F-4 SaR Dec 2005 ADF Ambiguity Page 111 Part-66 Goniometer Principle In the first generation of ADF systems, a real loop was rotated to find out the direction to the ground station. In the next generation of receivers - which is still on duty - , the loop antenna has converted to a Goniometer inside the receiver itself. The so called loop antenna now has a sine and a cosine coil fixed to the aircraft structure and the amplitudes and phases of the received signals in these coils determine the direction where the signal comes from. In the latest generation of ADF systems no mechanical moving parts are used to determine the direction. All the calculations are done by microprocessors. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 112 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 FERRITE SINE/COSINE ANTENNAS BEARING INDICATION FIELD COILS For Training Purposes Only SEARCH COIL RECEIVER MOTOR Figure 57 HAM US/F-4 SaR Dec 2005 ADF-Goniometer principle Page 113 Part-66 Quadrantal Error Each metal object of appropriate size functions as an antenna, and each receiving antenna retransmitts half of the received power. Also the fuselage of the aircraft works as antenna and as a result, the ADF bearings are shifted towards the logitudinal axis of the aircraft. Especially bearings around the 450 angle to the fuselage are bended. This is called the Quadrantal Error. It is compensated by a Quadrantal Error Corrector (QEC). The QEC simply adjusts the output of the Cosine coil to the output of the Sine coil with a voltage devider. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 114 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 58 HAM US/F-4 SaR Dec 2005 Quadrantal Error Page 115 Part-66 AIRCRAFT EQUIPMENT The aircraft is equipped with one or two ADF system. Each system consists of: S ADF Control Panel S Radio Magnetic indicator S ADF Antennas ( LOOP/ SENSE ) S ADF Receiver In the system the ADF receiver is connected to: − the Control Unit ( ADF Cont. PNL or MCDU or RMP ) − the loop and sense antenna ( combined ) − the RMIs or NDs via EFIS SGUs − the Audio Integrated System . The control unit provides the frequency command for the receiver, and the switching of function modes. The antennas contain the loop and sense elements and 3 buffer amplifiers. The loop element may include a testing coil, so that an ADF test signal can be fed into the antenna for test purpose. The transmission lines from the antenna to the receiver have an impedance of 78 Ohm. The receiver data output is connected to the RMIs and the NDs (via EFIS SGUs) to indicate the bearing information. The receiver audio outputs supply the audio integrated system with the related identification signals or the voice signals from the ground station. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 116 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Part-66 ADF CONTROL PANEL or MCDU or RMP ADF TEST BFO ANT OFF ADF LOOP ANTENNA ADF RMI QEC EFIS SGU RELATIVE BEARING For Training Purposes Only ANT CPLR SENSE ANTENNA ADF on ND ADF - RECEIVER ADF RMI Figure 59 HAM US/F-4 SaR Dec 2005 ADF Blockdiagram Page 117 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION ADF Part-66 ADF CONTROL PANEL The control panel allows to select the frequency and some modes like: S ADF S ANT S LOOP (old airplanes) S BFO S TEST Volume control (if installed) adjusts loudness of voice and identification. Features are: S Two liquid crystal frequency display windows. The frequency range of the unit is 190 thru 1750.5 kHz with 0.5 kHz spacing. S Two frequency selectors with three concentric knobs each. The inner knob controls the fractions and units of kHz, the middle knob controls the tens of kHz and the outer knob controls the hundreds of kHz. S A transfer (TFR) switch which is used to select one of two selected frequencies to tune the receiver. S Two green transfer lights which indicate the selected frequency used to tune the receiver. S A BFO/OFF switch which enables the reception of unmodulated A0/A1 signals in the BFO position or modulated signals in the OFF position. S An ANT/ADF switch which can be used if only reception of audio without direction finding is required (ANT position). S The processor in the unit converts the selected frequency into ARINC 429 format for transmission to the ADF receiver and supplies selected frequency data to drive the liquid crystal displays. HAM US/F-4 SaR Dec 2005 Page 118 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 No. 2 ADF Pointer No. 1 ADF Pointer ADF For Training Purposes Only VOL TEST BFO ANT OFF ADF Figure 60 HAM US/F-4 SaR Dec 2005 1 VOLUME CONTROL controls the audio output level 2. FREQUENCY INDICATOR 3. TEST SWITCH ADF needle moves to a distinct position 4. ADF/ANT SWITCH selects the operating mode 5. FREQUENCY SELECTOR 6. BFO (beat frequency oscillator) provides a tone allowing unmodulated CW keying to be heard ADF Control Panel Page 119 Part-66 AUDIO If ADF is selected on the Audio Control Panel, the 1020 Hz identification Morse code can be heard in the loudspeaker. A selector switch (if installed) allows to select identification only (RANGE), BOTH and VOICE only. NDB / Locator Transmissions and Audio For bearing information, it is only necessary for the ground station to transmit the carrier wave continously. But about every 30 seconds, the two or three letter identification Morse code is transmitted. This can happen as a keyed carrier only (A0/A1-modulation) or with a tone modulation of 1020 Hz (A0/A2-modulation). In some cases also speach modulation may be possible. Normaly both speach and ident tone can be heard in the audio system at the same time. To separate them, it is possible to switch a filters for the 1020 Hz. If only the 1020 Hz are passed and the voice is canceled, the switch position is called RANGE or IDENT. If only the 1020 Hz is canceled, the switch position is called VOICE ONLY. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 120 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 61 HAM US/F-4 SaR Dec 2005 Ident and Voice Frequency Range Page 121 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Part-66 ADF INDICATIONS Before showing the indicators, some terms in navigation had to be clarified. So called Q-codes are used in Morse-telegraphy. Four of them are discussed in the following. S QDM is the magnetic bearing to a ground station S QDR is the magnetic bearing from a GND station S QTE is the true bearing from a GND station S QUJ is the true bearing to a ground station Mathematical relations between the terms are: QDM = MH + RB QDR = QDM + 1800 QTE = QUJ + 1800 For Training Purposes Only magnetic true FROM station to aircraft QDR QTE from aircraft TO station QDM QUJ HAM US/F-4 SaR Dec 2005 Page 122 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Part-66 For Training Purposes Only NDB Figure 62 HAM US/F-4 SaR Dec 2005 Relative Bearing RB Page 123 Part-66 Radio Magnetic Indicator The pointers peak shows QDM, the pointers tail shows QDR. If no reception is possible, the pointer moves in a horizontal position. The test indication is defined by the manufacturer and may be for example 450 left of the lubber line. In case of receiver or indicator failure, a red flag may come in to view. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 124 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Part-66 For Training Purposes Only V O R V O R ADF ADF Figure 63 HAM US/F-4 SaR Dec 2005 ADF RMIs Page 125 Part-66 COMPONENTS LOCATIONS S Loop Antenna S Sense Antenna at the bottom of the fuselage sometimes extra, nowadays integrated in the loop antenna case − Note: in modern A/C, both antennas are normally in one case and located at the top of the fuselage S ADF Receiver in the electronic compartment S ADF Control Panel in cockpit S Radio Magnetic Indicator RMI or EHSI / ND in cockpit, pilots panels For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 126 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 For Training Purposes Only modern aircraft antenna installation Figure 64 HAM US/F-4 SaR Dec 2005 ADF Components Locations Page 127 Part-66 ADF RECEIVER SCHEMATIC Tuning Tuning range is from 190−1749.5 KHz. The control panel tuning mechanism is a Binary Coded Decimal System (BCD) as follows: A = 20 = 1 B = 21 = 2 C = 22 = 4 D = 23 = 8 The largest knob selects thousands (A only) and hundreds of kilohertz (A−B− C−D). The middle knob selects tens of kilohertz (A−B−C−D). The small knob selects units (A−B−C−D) and .5 KHz (A only) The coded output of the ADF control panel, is routed to the frequency synthesizer and is used to tune the band logic and hence the RF loop amplifier and band pass filters. Also the synthesizer output is mixed with the 1st and 2nd mixers as well as the self test mixer. The audio is muted during the tuning process and the AGC amplifiers are disabled. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 ANT MODE Upon selection of the ant mode a ground is applied from pin C of the control panel to pin 37 of the receiver and sensed at the input of the level detectors and converted to plus 5V DC. The 5V DC is used as a logic 1 and is routed to AND gate 1. If the self test button is not pressed, OR gate 2 output is a logic 1, disabling the loop RF amplifier and enabling the horizontal positioner. The horizontal positioner uses X−Y voltages from the output synchro transmitter as an error input. The error input is summed with a 26V AC reference and is routed out of the positioner to the servo phase detector via the servo amplifier. The motor is driven in a direction, dependent on the phase of the error signal and positions the synchro transmitter to null out the error. When a null is reached, the positions of the RMI ADF No. 1 pointer will be at 9 o’clock (or horizontal with reference to the RMI case). This is the stowed position of the RMI pointer when antenna mode is selected. The ground station signal is received by the sense antenna and is connected via the ADF sense coupler to the sense tuned circuit and RF sense amplifier. The RF amplifier output is routed,through the band pass filters to the 1st mixer. The signal is mixed with a frequency synthesizer RF signal to develop a 15 MHz Ist IF signal. The IF is amplified and again mixed with a second frequency synthesizer RF signal to develop a 3.6 MHz 2nd IF signal. The second IF signal is amplified, detected and routed to the AGC circuit to develop gain control of the 3.6 MHz and 15 MHz IF amplifiers. The output of the AM detector is routed to an audio amplifier and transformer. The audio output amplifier is varied by the gain potentiometer in the control panel and then routed to the audio integrating shelf. Selection of tone by the tone−off switch enables the 1020 Hz tone generator. The 1020 Hz tone is sent to the 3.6 MHz filter and is used to receive CW identification signals. Page 128 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Figure 65 HAM US/F-4 SaR Dec 2005 ADF Receiver Page 129 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ADF Part-66 ADF Mode In the ADF mode, the loop signal is sensed by the two loops in the loop antenna and routed to the receiver via the QEC. The goniometer, in the receiver, resolves the lateral and longitudinal loop signals into a single voltage. This RF signal is connected to a loop transformer and 900 phase shift circuit located in the RF amplifier. The signal then connects to a balanced modulator whose modulation frequency is controlled by a 90 Hz oscillator. The output of the balanced modulator is connected to a summing junction where the modulated loop signal is combined with the sense antenna signal. The received sense antenna signal is 900 out of phase with the received signal of the loop antenna. Since the loop signal is shifted by 900 in the phase shift circuit, the sense antenna signal arrives at the summing junction either in phase or 1800 out of phase with the loop antenna signal. For a given airplane position, an in phase condition will cause the RMI pointer to be driven in one direction and an out of phase condition will cause the RMI pointer to be driven in the other direction. Since all directional antennas have two null points these circuits are necessary to cause the servo system to always be driven to the correct null. The resultant phase modulated signal is routed to the 1st mixer via the bandpass filters. The 1st mixer combines this signal with a local oscillator signal generated by the frequency synthesizer. The resultant 15 MHz IF signal is filtered, amplified, and routed to the 2nd mixer. The 2nd mixer mixes the 1st IF with the synthesizer’s local oscillator signal to develop a 3.6 MHz 2nd IF signal. This signal is filtered, amplified, and sent to a phase modulation detector and an amplitude modulation AM detector. If the signal is of adequate strength an output of the phase modulation detector represents an error signal to the servo amplifier and servo phase detector. The detector’s output is amplified and used to drive the servo motor which in turn drives the goniometer to null out any existing error signal. When a null is reached the motor will also have driven the Receiver Synchro Transmitter and the RMI pointers will be positioned to indicate relative bearing to a selected ground station. The AM detected signal develops an AGC voltage to control the 1st and 2nd IF amplifiers and also goes to a signal level detector and to an audio amplifier. The audio level is varied by a gain control in the control panel before being routed to the audio integration circuits. If the amplitude modulation signal is lost, the signal level detector sends a logic 1 or OR gate 2. This causes the OR gate to enable the 9 o’clock horizontal positioner and disable the loop RF amplifier. As in the ant mode the horizontal positioner causes the RMI pointer to be driven to the 9 o’clock position. HAM US/F-4 SaR Dec 2005 Self Test Self test of the ADF system can be performed from either the receiver or the control panel. The control panel test switch is a momentary. switch. When the test switch is pressed, a ground is applied to the receiver self test circuits. The ground disables AND gate 1 and also energizes a test relay which removes the loop antenna signal from the circuit. The ground is also inverted to a “one” level and is routed to OR gate 3 turning on the 1020 Hz tone generator and to a switch that shorts out the sense antenna receiver input. Also the logic 1 enables the self test 15 MHz oscillator and mixer. The output of the mixer is routed to the goniometer as two equal inputs and to the RF sense amplifier. These test signals combine ahead of the bandpass filters and cause the test indicator, on the receiver, to indicate 450 and the RMI 1 pointer on the indicator to drive to a relative bearing of 3150 50, referenced to the RMI case. With the tone generator enabled, the 1020 Hz tone will be heard as long as the mode switch is held in the test position. Page 130 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ADF Figure 66 HAM US/F-4 SaR Dec 2005 ADF Receiver Page 131 Part-66 APPENDIX For the RMI it is necessary to get the heading information to drive the compass rose. This heading information can be supplied from either: S Directional gyro (DG) slaved by a flux valve S Attitude Heading Reference System (AHRS) S Inertial Navigation system (INS) S Inertial reference system (IRS, ADIRS) For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 132 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 67 HAM US/F-4 SaR Dec 2005 Compass Information for the RMIs Page 133 Part-66 Modern Aircraft Equipment Indication of ADF bearings are integrated in the EFIS system, if the aircraft is equipped with electronic cathode ray tube displays (CRT). One advantage is, that the indication of the bearing pointers will automatically disappear, if a failure in the ADF system is detected. Additionally, a flag warning may be displayed. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 134 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 68 HAM US/F-4 SaR Dec 2005 EFIS Page 135 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 69 HAM US/F-4 SaR Dec 2005 ND - ADF Pointers Page 136 Part-66 LEGEND 1 = ADF Ground Station Location and Identification D = Bearing Pointer ADF 1 − Single Pointer ADF 2 − Double Pointer 3 = Tuned ADF Station Data 4 = ADF/VOR Selector Switches 5 = ADF−1 Pointer 6 = No. 2 Pointer, set to VOR Operation in this example 7 = Warning Flags (indicating a Failure in ADF−1 and VOR−2) 8 = The Bearing Pointers are biased to 3 o clock posn. with a system failure. 9 = ADF Failure Indication on ND (Note: The pointer is biased out−of−view). For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 70 HAM US/F-4 SaR Dec 2005 ADF - Data and Warning Displays Page 137 Part-66 Navigation Control S Normal tuning NAV radios are normally controlled by the FMGEC. Normal VOR/DME, ILS and ADF tuning is automatic in accordance with the procedures selected by the crew via the FMGEC MCDU. S Manual tuning The VOR, DME, ILS and ADF frequencies may be changed as required, using the FMGEC MCDU RADIO NAV page (see below). The ADF BFO is controlled using a line key on this page to enable identification of A1 modulated NDB’S. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF HAM US/F-4 SaR Dec 2005 Page 138 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 71 HAM US/F-4 SaR Dec 2005 Navigation Frequency Tuning Page 139 Lufthansa Technical Training M13.4 NAVIGATION ADF Part-66 Radio Navigation Back-Up In normal configuration each FMGEC controls its on-side receivers. In case of one FMGEC failure, the remaining one controls both sides receivers For Training Purposes Only In case of both FMGEC’s failure, a navigation back-up is provided by two RMP’s located on the pedestal, each controlling its on-side receivers. HAM US/F-4 SaR Dec 2005 Page 140 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ADF Figure 72 HAM US/F-4 SaR Dec 2005 Radio Navigation Back-Up Page 141 Lufthansa Technical Training EGPWS For Training Purposes Only Additionally, Windshear alerting ( mode 7 ) can be provided for specific aircraft types. M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 GENERAL The Enhanced Ground Proximity Warning System incorporates the functions of the basic GPWS. The basic GPWS is operative over a range from 30 ft to 2450 ft. radio altitude and includes the following alerting modes ( modes 1 to 6 ): HAM US/F-4 SaR Dec 2005 The EGPWS has all the features of the GPWS basic modes and the following additional functions: S Terrain Awareness Alerting and Display. S Terrain Clearance Floor, The purpose of the Enhanced Ground Proximity Warning System (EGPWS) is to help prevent accidents caused by Controlled Flight into Terrain (CFIT). This is done by comparing the aircraft position ( FMS or GPS ) to an internal database. The system achieves this objective by accepting a variety of aircraft parameters as inputs, applying alerting algorithms, and providing the flight crew with aural alert messags and visual annunciations and displays in the event that the boundaries of any alerting envelope are exceeded. Page 142 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 73 HAM US/F-4 SaR Dec 2005 Type of Accident Page 143 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 EGPWS EVOLUTION The first generation of GPWS, MARK I, was developed in the early 1970‘s because of a FAA ruling in 1974. Continuous improvements have been made to the GPWS as a result of this experience, and two new products were introduced in the late 1980‘s: S MARK V for digital aircraft and S MARK VII for analog aircraft. First generation systems do not give the pilot information about the cause of the warning. Warnings in this situation can be perceived as “false“. The voice warning “Too Low Gear“ clearly identifies the problem. The implementation of GPWS has been very effective at reducing CFIT accidents. About 95% of the world‘s fleet is equipped with GPWS, but CFIT continues to be the number one cause of airline passenger fatalities worldwide. Much of this continuing problem is due to CFIT situations not addressed by early generation GPWS. Today, a very large portion of CFIT accidents occurs in the landing configuration on a stable non−precision approach, in low visibility, but short of the runway. In this configuration, the GPWS expects the airplane to land and will not issue a warning. Nuisance warnings have been problematic in early generation GPWS and, in fact, have caused pilots to ignore valid warnings in CFIT accidents. Through the years, much of the GPWS development and improvements have been focused on eliminating nuisance warnings. Terrain clearances at certain airports can also cause nuisance or short warnings. Modern AlliedSignal GPWS have incorporated Envelope Modulation that monitors aircraft position relative to a database of specific troublesome terrain−airport locations. When the GPWS recognizes that the aircraft is landing at one of these airports, it will adjust the warning modes for that airport, but only under a specific state of conditions. The evolution of GPWS through the years has been a history of steady and continuous improvement. While the data shows a marked reduction in CFIT, primarily due to GPWS, data also shows that CFIT remains the number one safety concern. The next significant improvement to GPWS is the Enhanced GPWS (EGPWS). EGPWS was developed to address the remaining CFIT issues and to add virtual look−ahead capability. A recent study of CFIT accidents analyzed the GPWS performance in each case with the results shown in Figure 74 : HAM US/F-4 SaR Dec 2005 EGPWS improvements were focused on solving these problems: S No Warning: The primary cause of CFIT is landing short with no GPWS warning. When the landing gear is down and landing flaps are deployed, the GPWS expects the airplane to land and therefore, issues no warning. EGPWS introduces the Terrain Clearance Floor (TCF) function, which provides GPWS protection in the landing configuration. S Late Warning or Improper Response: The occurrence of a GPWS alert typically happens at a time of high workload and nearly always surprises the flight crew. Almost certainly, the aircraft is not where the pilot thinks it should be, and the response to a GPWS warning can be late in these circumstances. Warning time can also be short if the aircraft is flying into steep terrain since the downward looking radio altimeter is the primary sensor used for the warning calculation. The EGPWS improves terrain awareness and warning times by introducing the Terrain Display and the Terrain Data Base Look Ahead protection. EGPWS builds on the proven success of the GPWS and continues improvement focusing on the actual causes of CFIT. A comparison of older generation systems with the EGPWS is shown in Figure 74 : The Mark VI EGPWS utilizes regions of a global Terrain Database that is organized in a flexible and expandable manner. Using digital compression techniques, the complete database is stored in non−volatile memory within the LRU. Updates and additions are easily done via the LRUs front−panel and a PCMCIA card. Page 144 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 COMMERCIAL JET AIRCRAFT ACCIDENTS 1988−1995 Late Warning or Improper Pilot Response 41% No GPWS installed 31% No Warning 28% Comparison of GPWS Generations For Training Purposes Only MARK I MARK II/III MARKV/VII EGPWS BASIC GPWS WITH 5 MODES X X X X ENUNCIATE CAUSE OF WRNG X X X AIRSPEED MODE EXPANSION X X X ALTITUDE CALL−OUTS X X BANK ANGLE WARNING X X ENVELOPE MODULATION X X TERRAIN CLEARANCE FLOOR X LOOK−AHEAD WARNING X TERRAIN DISPLAY X NUISANCE WARNINGS OFTEN COMMON RARE RARE Figure 74 HAM US/F-4 SaR Dec 2005 Aircraft Accidents Statistic Page 145 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 OPERATION The EGPWC knows 6 basic modes of ground proximity warning. The modes 1 through 5 are basic system requirements. Mode 6 provides additional protection in form of a selectable menu of radio altitude callouts during the landing approach. It also provides an optional alert for excessive bank angles. The modes are program−pin selectable before installation of the EGPWC. − interface with the CMC/CFDS maintenance systems with interactive protocols when the aircraft is on the ground, − front panel PCMCIA operations to upload software and databases, − front panel maintenance test operations for system checkout and troubleshooting, − front panel system status readouts to monitor system conditions. An overall EGPWS block diagram of the system also shows the enhanced additional components. The Enhanced Ground Proximity Warning Computer (GPWC) is the control unit of the EGPWS. The primary functions of the computer are: S Receiving and processing of serial digital and discrete data for use in computing warning mode conditions. S Warning mode computations to determine when the aircraft penetrates one of the warning boundaries mentioned above. S Output warnings and control S audible and visible indications when a mode computation reveals a boundary penetration. S Monitoring and indicating of the GPWS status as well as internal monitoring of the GPWC. The computer stores failures for later readout on a front panel BITE display. The main part of the EGPWS operations is the warning function. This function is independent of the other functions. For example, the loss of the terrain awareness display function will not affect the operation of the ground proximity warning functions as long as the necessary inputs are still available. In addition to the main alerting functions, the EGPWC also performs the subsequent auxiliary functions: − input signal processing, to include filtering and signal monitoring, − alert output processing, to include alert prioritization, voice message synthesis, audio output and display, warning lamp drivers, − Built−In Test Equipment (BITE) and monitor, to include cockpit−activated self−test, HAM US/F-4 SaR Dec 2005 Page 146 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 For Training Purposes Only Headset Figure 75 HAM US/F-4 SaR Dec 2005 EGPWS Simplified Schematic Page 147 Part-66 MODE 1 − Excessive Descent Rate When the aircraft penetrates the outer alert boundary the EGPWS generates the aural message ”SINKRATE”. When the aircraft penetrates the inner alert boundary the aural message ”PULL UP” is generated. In both cases the computer also generates a discrete output to drive visual annunciations. The computer defines the two alert boundaries by inputs of barometric vertical speed, inertial vertical speed and radio altitude. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 148 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 76 HAM US/F-4 SaR Dec 2005 Mode 1 Excessive Descent Rate Page 149 Part-66 MODE 2 − Excessive Terrain Closure Rate When the aircraft closes in on the terrain with a hazardous rate the EGPWS generates MODE 2 alerts. For MODE 2 alerts the aircraft can be in a S descent, S level flight or in a S climb. The system uses a combination of vertical speed and radio altitude signals when it computes the mode alert boundaries. There are two submodes, MODE 2A and MODE 2B, determined by aircraft configuration. − MODE 2A The MODE 2A alert is generated when the aircraft is in any configuration other than defined for MODE 2B. When the aircraft penetrates this mode’s alerting boundaries the computer generates the aural message ”TERRAIN TERRAIN” together with visual annunciation signals. If the aircraft continues to penetrate the boundaries, the aural message changes and ”PULL UP” is repeated until the aircraft exits the warning envelope. The upper alert boundary varies as a function of aircraft speed. − MODE 2B The MODE 2B alert is generated when the aircraft is in the subsequent configuration: S the flaps are set to the landing position, S the aircraft position is within +/−2 dots of the localizer and/or the glide slope centerline indication during an ILS approach, S for the first 60 seconds after takeoff. If the aircraft penetrates the MOD 2B envelope with either gear or flaps not in the landing configuration, the aural message ”TERRAIN TERRAIN” is generated initially. If the aircraft continues to penetrate the envelope, the aural message ”PULL UP” is repeated until the aircraft exits the envelope. In both cases the computer also sends signals to drive visual annunciators. If both gear and flaps are in landing position the aural message ”TERRAIN” is repeated until the aircraft leaves the warning envelope. The aural „ PULL UP „ messages are suppressed. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 150 Part-66 MODE 2 A MODE 2 B For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 77 HAM US/F-4 SaR Dec 2005 Mode 2 Excessive Terrain Closure Rate Page 151 Part-66 MODE 3 − Altitude Loss After Takeoff The MODE 3 function provides alerts when the aircraft looses a significant amount of altitude immediately after takeoff or during a missed approach. The EGPWC receives altitude above sea level information. When a loss of altitude is detected the computer calculates the amount of permitted loss before an alert is initiated. When the landing gear is up and the flaps are not in landing position after takeoff or a go around, the computer initiates the MODE 3 alert, if an altitude loss larger than permitted is detected. This alert will remain on until the aircraft reaches sufficient altitude and is no longer in the takeoff phase of the flight. The MODE 3 warning is an aural message ”DON’T SINK” and an activation of visual annunciators. The visual annunciators remain active until a positive rate of climb is established. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 152 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 78 HAM US/F-4 SaR Dec 2005 Mode 3 Altitude Loss After Takeoff Page 153 Part-66 MODE 4 − Unsafe Terrain Clearance The EGPWS provides MODE 4 alerts and warnings when insufficient terrain clearance is detected in reference to phase of flight and speed. There are 3 submodes. − MODE 4A The MODE 4A is active during cruise and approach with the landing gear not in landing configuration. The standard upper boundary is 500 feet radio altitude. If the aircraft penetrates this boundary from above with the landing gear still up at 190 knots or below, the aural message ”TOO LOW GEAR” is generated. The upper boundary increases with airspeed linearly. It is at its maximum of 1000 feet radio altitude at 250 knots and above. If this boundary is penetrated, the aural message ”TOO LOW TERRAIN” is generated. − MODE 4B For the MODE 4B function the upper boundary decreases to 245 feet radio altitude when the landing gear is lowered. Penetration of the boundary at speeds below 159 knots results in a ”TOO LOW GEAR” aural message with the gear up. With the gear down but the flaps not in landing configuration the aural message ”TOO LOW FLAPS” is generated. The EGPWC generates the aural message ”TOO LOW TERRAIN” at speeds above 159 knots, if the boundary is penetrated. − MODE 4C MODE 4C is based on a minimum terrain clearance or floor that increases with radio altitude during takeoff. A value equal to 75% of the current radio altitude is accumulated in a long term filter. Any decrease of radio altitude below the filter value with gear and flaps up results in the aural message ”TOO LOW TERRAIN”. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 154 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 MODE 4 A ( GEAR UP ) MODE 4 C ( AT TAKEOFF ) MODE 4 B For Training Purposes Only ( GEAR DOWN ) Figure 79 HAM US/F-4 SaR Dec 2005 Mode 4 Unsafe Terrain Clearance Page 155 Part-66 MODE 5 − Below Glide Slope The EGPWC provides 2 levels of MODE 5 operations when the aircraft flight path deviates below the glide slope beam during a front course ILS approach. The ”soft” alert activation occurs whenever the aircraft is more than 1.3 dots below the beam center. The volume level of the aural alert message ”GLIDE SLOPE” is approximately one half (−6dB) of the other alert message volume levels. The ”hard” alert activation occurs below 300 feet radio altitude whenever the aircraft is more than 2 dots below the glide slope. The aural alert message ”GLIDE SLOPE” is generated at full normal volume level. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 156 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 80 HAM US/F-4 SaR Dec 2005 Mode 5 Excessive Glide Slope Deviation Page 157 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 MODE 6 − Call−Outs The EGPWS provides MODE 6 alerts and call−outs when the aircraft descends below predefined altitudes, Decision Height (DH), minimums, approach decision height or approaching minimums. This mode also generates alerts and warnings for excessive roll or bank angles. The operator can select specific call−outs by program pins from predefined menus. Modes 6 alerts and call−outs generate aural messages and ARINC 429 outputs, but no visual indications. Bank Angle Mode 6 bank angle callouts occur when the airplanes bank angle is larger than 10 degrees between 30 feet and 130 feet. Above 130 feet the callout occurs at 35 degrees, 40 degrees, and 45 degrees. The aural message is “BANK ANGLE, BANK ANGLE“. For Training Purposes Only Altitude callouts Altitude callouts start at 2500 feet. At 2500 feet, there is an option to give the aural ”TWENTY FIVE HUNDRED” or give the aural ”RADIO ALTITUDE”. The minimums callout option gives an aural callout when the airplane descends through the decision height altitude set on the EFIS control panels. These are the aural callouts the GPWC can give for decision height: − Minimums − Minimums minimums − Decision height. There is also an approaching minimums option callout that tell the pilots when the airplane is approaching the decision height set on the EFIS control panel. The callout normally comes on when the airplane altitude is 80 feet above the decision height. The aural callouts for this option are: − Approaching minimums − Approaching decision height − Plus hundred (for this callout the altitude selection is set for decision height + 100 feet). HAM US/F-4 SaR Dec 2005 Page 158 Part-66 Altitude Call-Outs Excessive Bank Angle For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 81 HAM US/F-4 SaR Dec 2005 Mode 6 Call-Outs Page 159 Part-66 MODE 7 − Windshear Alerting (TSO−C117A) Mode 7 produces optional alerts for flight into an excessive Windshear conditions during takeoff or final approach in accordance with TSO−C117. The Windshear Caution, or pre−alert as it is sometimes termed, provides visual and ARINC 429 output indications. The Windshear warning also produces aural, visual and ARINC 429 output indications. This mode is selected with a program pin strap. Windshear detection is active between 10 and 1500 feet AGL during the initial takeoff and final approach phases of flight. Alerts and warnings are provided when the level of windshear exceeds predetermined threshold values. The actual windshear value measured represents the vector sum of inertial vs. air mass accelerations along the flight path and perpendicular to the flight path. These shears result from vertical winds and rapidly changing horizontal winds. Windshear warnings are given for: S decreasing head wind (or increasing tail wind) and severe vertical down drafts. S increasing head wind (or decreasing tail wind) and severe up drafts. The windshear microburst phenomenon and windshear caution and warning levels are illustrated in Figure 82 . For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 160 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 82 HAM US/F-4 SaR Dec 2005 Mode 7 Excessive Windshear Detection ( RWS ) Page 161 Part-66 ENVELOPE MODULATION The envelope modulation feature of the EGPWS provides alerting protection at some key locations throughout the world. At other locations nuisance margins are improved. This is made possible with the use of navigational signals from modern area navigation equipment. This feature optionally utilizes the Global Positioning System (GPS) or updated flight management system navigational signals, if available. The airports, that need envelope modulation, are in the GPWC memory and identified by their latitude and longitude. When the airplane is near one of these airports, the GPWC looks at other airplane parameters to make sure an envelope modulation condition exists. The additional parameters are: − Radio altitude − Magnetic track − Glide slope deviation − Localizer deviation − Selected runway heading − Barometric corrected altitude. Modes 4, 5 and 6 are expanded at some locations to provide alerting protection consistent with normal approaches. Modes 1, 2 and 4 are desensitized at other locations to prevent nuisance warnings that result from unusual terrain or approach procedures. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 162 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 83 HAM US/F-4 SaR Dec 2005 Envelope Modulation Page 163 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 TERRAIN AWARENESS ALERTING AND DISPLAY A major new feature of the EGPWS is the incorporation of the Terrain Awareness Alerting and Display functions. These functions use the aircrafts geographic position and altitude and the internal terrain database to predict potential conflicts between the aircraft flight path and the terrain and optionally provide graphic displays of the conflicting terrain. The feature is illustrated by the block diagram in Figure 84 . Terrain Awareness Inputs The EGPWC receives airplane data from the air data inertial reference system (ADIRS) and the global positioning system (GPS). The terrain awareness function uses this data: − Latitude − Longitude − Barometric altitude − Ground track − Ground speed − Heading − Roll attitude − Flight path angle (calculated by GPWC). Terrain awareness uses GPS or FMS for latitude and longitude. For Training Purposes Only Terrain Awareness Calculation The EGPWC has a world−wide terrain database in memory. The EGPWC looks at airplane position and track and compares this data to the terrain database. If the EGPWC finds there is a terrain threat, it makes an alert. Terrain Display Output The EGPWC makes a digital map of the terrain forward of the airplane. It sends this digital map to the display units (DUs) to show on the navigational displays (NDs). The display uses different colored dots to show terrain altitude relative to airplane altitude. HAM US/F-4 SaR Dec 2005 Page 164 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 AMU GPS PFD For Training Purposes Only ADIRS FMS DMC / DIU EGPWS Figure 84 HAM US/F-4 SaR Dec 2005 ND Terrain Display Page 165 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 TERRAIN ALERTING Two alerting envelopes are computed, one corresponding to a S caution-level alert and the other to a S warning-level alert. S Caution The look-ahead caution alert is provided approximately 40 to 60 seconds before a potential terrain conflict. The caution level aural message is ”CAUTION TERRAIN” and repeated every seven seconds while within the terrain caution envelope. Alert discretes are provided to energize the terrain display relays. S Warning The look-ahead warning alert is provided approximately 20 to 30 seconds befor a potential terrain conflict. The warning level aural message is ”TERRAIN TERRAIN PULL UP” and ”PULL UP” is repeated continuously while within the terrain warning envelope. Alert discretes are provided to energize the terrain display relays. The alerting envelopes are sized by a look−ahead distance in front of the airplane, an altitude offset below the airplane, and a lateral distance on either side of the airplane (Figure 85 ). The look−ahead distance varies mainly with ground speed; as ground speed increases, the alerting distance increases to provide roughly equivalent alerting times at all speeds. The look−ahead distance is primarily focused along the airplane flight path (climbing, descending or level). An additional component looks six degrees up to protect against very high terrain (shown in Figures on the right ). This six−degree component actually looks ahead approximately twice the normal look−ahead distance. The altitude offset is normally 700 feet (213m) below the airplane. The purpose of the offset is to provide terrain alerting when the airplane has less than the normal terrain clearance. The lateral distance is 1/8 nautical mile (0.2 3 km) either side of the airplane ground track and increases gradually out to the look-ahead distance. To enable the airplane to land without nuisance alerts, the look−ahead distance and altitude offset decrease as the airplane approaches the airport. TERRAIN CAUTION & WARNING ENVELOPE BOUNDARIES SLOPES=GREATER OF FPA OR +6 FLIGHT PATH ANGLE (FPA) TERRAIN FLOOR WARNING AREA CAUTION AREA SLOPES VARY WITH FPA WARNING LOOK AHEAD DIST CAUTION LOOK AHEAD DIST WARNING LOOK UP DIST CAUTION LOOK UP DIST PERSPECTIVE VIEW OUTSIDE LINES POINT OUT " 3 CENTER LINE POINTS ALONG GROUND TRACK PLUS A LEAD ANGLE DURING TURNS STARTING WIDTH=1/4 nM LOOK AHEAD DISTANCE HAM US/F-4 SaR Dec 2005 Page 166 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 Caution envelope Look-ahead distance Altitude offset Example: Caution alert at 2 nM ( 45 sec @ 180 Kn ) Warning envelope For Training Purposes Only Look-ahead distance Altitude offset Example: Warning alert at 1.1 nM ( 22 sec @ 180 Kn ) jc Figure 85 HAM US/F-4 SaR Dec 2005 Terrain Alerting Page 167 Lufthansa Technical Training M13.4 NAVIGATION EGPWS Part-66 TERRAIN DISPLAYS AND ALERTS The terrain awareness alerting and display function maintains a background display of local terrain forward of the aircraft for cockpit display. In the event of terrain caution or warning conditions, an aural alert and lamp outputs are triggered. The background image is then enhanced to highlight related terrain threats forward of the aircraft. NOTE: TERRAIN IS NOT SHOWN IF MORE THAN 2000 FT BELOW REFERENCE ALTITUDE AND / OR TERRAIN IS NOT SHOWN IF TERRAIN ELEVATION IS WITHIN 400 FT OF RUNWAY ELEVATION NEAREST THE AIRCRAFT. For Training Purposes Only Color Threat Solid Red Warning terrain (approx. 30 sec. from impact); audio alert TERRAIN AHEAD, PULL UP Solid Yellow Caution terrain (approx. 60 sec. from impact); audio alert: TERRAIN AHEAD High Density Red 50 Terrain that is more than 2000 ft. above reference altitude* High Dens. Yellow 50 Terrain that is between 1000 and 2000 ft. above reference altitude* Medium Dens. Yellow 25 Terrain that is 500 (250 with gear down) ft. below to 1000 ft. above reference altitude* Medium Dens. Green 25 Terrain that is 500 (250 with gear down) ft. below to 1000 ft. below reference altitude* Light Dens. Green 12.5 Terrain that is 1000 to 2000 ft. below reference altitude* Black No close terrain Light Density Magenta Unknown terrain * Reference altitude is aircraft altitude when climbing or during level flight, otherwise it is projected down from actual aircraft altitude to provide a 30 second advance display of terrain. HAM US/F-4 SaR Dec 2005 Terrain Caution Alert A specific audio alert and light output is triggered and the background image is enhanced to highlight the terrain caution threats. At the start of a terrain caution alert, the terrain awareness function triggers the caution audio alert phrase TERRAIN AHEAD. The phrase is repeated after seven seconds if still within the terrain caution envelope. During a terrain caution alert, the GPWS legend of pushbutton switches is on. During a terrain caution alert, areas where terrain violates the terrain caution envelope along the aircraft track, and within plus or minus 90 deg. of the aircraft track, are painted with the caution color 100 per cent yellow. Terrain Warning Alert When the conditions have been met to generate a terrain warning alert, a specific audio alert and light output is triggered and the background image is enhanced to highlight the terrain caution and warning threats. At the start of a terrain warning alert, the terrain awareness function triggers the warning audio alert phrase TERRAIN AHEAD, PULL UP. The phrase is repeated continuously while within the terrain warning envelope. During a terrain warning alert, the GPWS legend of pushbutton switches is on. During a terrain warning alert, areas where terrain violates the terrain warning envelope along the aircraft track, and within plus or minus 90 deg. of the aircraft track, are painted with the warning color 100 per cent red. NOTE: -WHEN AN ALERT OCCURS (CAUTION OR WARNING) AND THE FCU MODE IS NOT IN A CORRECT MODE (ARC OR ROSE), THE MESSAGE TERR. CHANGE MODE IS DISPLAYED ON ND’S. -WHEN AN ALERT OCCURS (CAUTION OR WARNING) AND THE FCU RANGE SELECTED IS 160 OR 320NM THE MESSAGE TERR. REDUCE RANGE IS DISPLAYED ON ND’S. TAD inhibitions S manually by the GPWS/TERR P.B. switch S automatically when the FMS aircraft position accuracy is not accurate enough. This is indicated to the crew by the automatic deselection of terrain display and the illumation of the TERR STBY memo on the ECAM display unit. Page 168 Lufthansa Technical Training M13.4 NAVIGATION EGPWS Part-66 TERR AHEAD -message TERR AHEAD in RED: Warning Terrain -message TERR AHEAD in AMBER: Caution Terrain For Training Purposes Only -message TERR in CYAN: normal indication Figure 86 HAM US/F-4 SaR Dec 2005 TERRAIN INDICATION ON ND Page 169 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 TERRAIN BACKGROUND −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− I FUNCTION COLOR/PATTERN −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− Solid Red Warning Terrain (approximately 30 seconds from impact) Solid Yellow Caution Terrain (approximately 60 seconds from impact) 50% Red Dots More than 2000 feet above reference altitude* 50% Yellow Dots 25% Yellow Dots 1000 to 2000 feet above reference altitude* 500 (250 with gear down) feet below to 1000 feet above reference altitude* 25% Green Dots 500 (250 with gear down) feet below to 1000 feet below reference altitude* 1000 to 2000 feet below reference altitude* 12.5% Green Dots For Training Purposes Only Black No close terrain Magenta Unknown Terrain * Reference altitude is aircraft altitude when climbing or during level flight, otherwise it is projected down from actual aircraft altitude to provide a 30 second advance display of terrain. HAM US/F-4 SaR Dec 2005 Page 170 Part-66 THIS PAGE INTENTIONALLY LEFT BLANK For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 171 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION EGPWS Part-66 Geometric Altitude Geometric Altitude is a computed aircraft altitude designed to help ensure optimal operation of the EGPWS Terrain Awareness and Display functions through all phases of flight and atmospheric conditions. Geometric Altitude uses an improved pressure altitude calculation, GPS Altitude, Radio Altitude, and Terrain and Runway elevation data to reduce or eliminate errors potentially induced in Corrected Barometric Altitude by temperature extremes, non−standard altitude conditions, and altimeter misssets. Geometric Altitude also allows continuous EGPWS operations in QFE environments without custom inputs or special operational procedures. Background Aircraft altitude, along with horizontal position and the terrain and obstacle database, are used to provide Terrain Awareness Display and Look−Ahead alerting and warning functions of the EGPWS. In order to ensure that timely alerts are provided and to minimize the potential for nuisance warnings, the EGPWS requires an accurate geometric altitude compatible with the elevation data contained in the terrain database. The primary source of altitude for aircraft operations is normally Corrected Barometric Altitude. Corrected Barometric Altitude is provided by an ADC using measured external static pressure and local pressure correction as entered by the flight crew. Corrected Barometric Altitude may be in error when compared to the true altitude, especially under extreme temperatures or non−standard atmospheric conditions such as inversion layers or strong pressure gradients. Corrected Altitude is also prone to errors induced by altimeter miss−sets and is not compatible with the EGPWS functions while operating under QFE conditions. Uncorrected Barometric (or Standard) Altitude is also available from the ADC. Uncorrected Altitude is not susceptible to QFE and altimeter miss−set errors but is not compensated for local pressure variations. An alternate source of altitude information is GPS, which provides a Geometric Altitude and is not significantly affected by atmospheric conditions. The overall accuracy of GPS Altitude however, is not typically sufficient to be used directly by the EGPWS, primarily due to errors induced by Selective Availability. However, GPS Altitude can be used in combination with other signals to provide a reliable estimate of its real time accuracy, which then can be used for reasonableness checking of other altitude sources. Since no single sensor can provide an accurate geometric altitude through all phases of flight and atmospheric conditions, the EGPWS computes an estimated average altitude using HAM US/F-4 SaR Dec 2005 S Pressure and GPS Altitudes, aircraft position, and the internal runway and terrain databases. This is the Geometric Altitude function. With the Geometric Altitude function, EGPWS can operate reliably throughout extreme local pressure or temperature variations from standard, is not susceptible to altimeter miss−sets by the flightcrew, and will not require any custom inputs or special procedures by the flightcrew when operating in a QFE environment. Page 172 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 87 HAM US/F-4 SaR Dec 2005 Terrain Background Display Page 173 Part-66 TERRAIN CLEARANCE FLOOR The Terrain Clearance Floor (TCF) alert function creates an increasing terrain clearance envelope around the intended airport runway directly related to the distance from the runway. TCF alerts are based on current aircraft location, nearest runway center point position and radio altitude. TCF is active: S during takeoff mode when Mode 4 protection is not available, and S during cruise and S during final approach. The EGPWC has a runway database in memory. This database contains the location of all hard surface runways in the world that are longer than 3,500 feet. TCF makes a terrain clearance envelope around the runway. The altitude of the envelope increases as the distance from the airport increases. EGPWC compares airplane latitude, longitude, and radio altitude with TCF envelope data. If the airplane descends through the floor of the envelope, GPWC makes an alert. ”TOO LOW TERRAIN” is enunicated upon penetration of the TCF. The Terrain Clearance Floor (TCF) alert message will occur two times when initial envelope penetration occurs, and one time thereafter for each 20% degradation in radio altitude. At the same time the GPWS warning lamps will illuminate. The lamps will remain on until the alert envelope is exited. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 174 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS Figure 88 HAM US/F-4 SaR Dec 2005 Terrain Clearance Floor Page 175 Lufthansa Technical Training M13.4 NAVIGATION EGPWS Part-66 TERRAIN-WEATHER INTERFACE Purpose The EGPWC and the WXR - Transceiver send display data that shows on the navigational displays (NDs). Two terrain weather select relays control which data shows on each ND. For Training Purposes Only Digital Relay Interfaces The EGPWC and WXR send data to both terrain weather select relays on ARINC 453 data buses. The terrain weather select relays send data to the display electronics units on ARINC 453 data buses. The display electronics units shows the data on the NDs. The EGPWS may be configured to automatically deselect the Weather Display and ‘pop−up‘ a display of the terrain threats when they occur. The logic for these controls accepts input ( HAZARD BUS ) from Predictive Windshear System (PWS), so that windshear alerts can override a terrain display and revert to the weather display with the corresponding windshear data. Relay Control When TERR is selected on either EFIS control panel , a terrain select ground discrete signal from the EGPWC allows the terrain / weather select relay to energize. This allows the terrain data to be connected to the DUs for display on the NDs. Select the TERR switch again and the relay deenergizes. The relay now connects the WXR to the DUs. If the Weather Radar system is ON, the XFR - relays remove the weather data bus from the EFIS signal generator, if the caution- or warning- criterias are fulfilled, and switch on to the terrain data bus. The navigation display unit will display the terrain areas which conflict with S the caution criteria in solid yellow or S the warning criteria in solid red color. The terrain switchlights will be illuminated. Subsequent pressing of the switchlight can deselect the terrain display and reselect the weather display. The background display is computed from the aircraft altitude with respect to the terrain data in the digital elevation matrix overlays. Where terrain data are available and sufficiently close to the aircraft altitude, they are presented in background color dot patterns reflecting the projected separations shown. Different dot density patterns and colors are used to represent terrain altitude bands with respect to the aircraft. Areas with no terrain data available are painted with the low density magenta. Known terrain sufficiently below the aircraft altitude is black. When weather radar is selected, the normal relay position lets the weather radar data show on the NDs. HAM US/F-4 SaR Dec 2005 Page 176 Lufthansa Technical Training M13.4 NAVIGATION EGPWS Part-66 CB Panel EGPWC For Training Purposes Only XFR Relay DMC/DIU 1 WXR HAZARD DMC/DIU 2 XFR Relay WXR R/T Figure 89 HAM US/F-4 SaR Dec 2005 Terrain / WXR Display Page 177 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION EGPWS Part-66 MKV EGPWS INTERFACE CONTROL PROGRAM PINS For the MKV EGPWS, various features and configurations are enabled through program pin strapping. Each program pin can be strapped to one of nine state pins, or left open. This provides 10 separate states for each program pin. The ten states are referred to as: STATE 1 = Program Pin open STATE 2 = Program Pin connected to PPCOM: MP3A STATE 3 = Program Pin connected to Discrete Output #1 (DO #1): MP15D STATE 4 = Program Pin connected to Discrete Output #2 (DO #2): MP3D STATE 5 = Program Pin connected to Discrete Output #3 (DO #3): MP7B STATE 6 = Program Pin connected to Discrete Output #4 (DO #4): TP12C STATE 7 = Program Pin connected to Discrete Output #5 (DO #5): TP12A STATE 8 = Program Pin connected to Discrete Output #7 (DO #7): TP4D STATE 9 = Program Pin connected to Discrete Output #6 (DO #6): TP12B STATE10= Program Pin connected to Discrete Output #10 (DO #10): TP11A Program pin monitoring is provided via a parity pin that is to be connected so that an odd number of program pins is strapped for any valid configuration.Program pins are read only during power up. Eight steps are required to perform a program pin read: one for each of the discrete outputs. The program pin configuration is read into NVM as long as there is not a Program Pin Read Error or a Program Pin Parity Error. S A Program Pin Read Error occurs if a pin toggles more then once during the read process, or a configuration change can not be confirmed with two consecutive program pin reads. The read process will be attempted 4 times before a read error is indicated. S A Program Pin Parity Error occurs when the program pin read process indicates that an even number of program pins have been connected. The program pins are always required to have odd parity. If either or both of these errors occur, the system will configure per the current valid NVM configuration. If no valid NVM configuration exists, then the system will configure to a zero default configuration (as if all pins open). The selected aircraft type will be enunciated as part of a self−test sequence ( Level 3 ). HAM US/F-4 SaR Dec 2005 Aircraft Type These 3 program pins allow for up to 512 (83) type selections. Appendix Figure 90 D−T1 lists some examples of the strapping details for each aircraft type and also includes cross−references to MKV aircraft type ID numbers. Page 178 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 90 HAM US/F-4 SaR Dec 2005 Aircraft Type Strapping Table D-T1 Page 179 Part-66 Audio Menu The 3 program pins allow for up to 512 (8 3 ) type selections. These menus are only active when the Callout Enable Discrete is active. If the discrete is not selected, then only the MINIMUMS−MINIMUMS alert function is enabled Appendix Figure 91 defines each of the Altitude callout menu selections. The selected callout menu will be enunciated as part of a self−test sequence. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 180 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 BASIC VOICE MENU ALERT/WARNING CONDITION MODE 7 WINDSHEAR WARNING MODE 6 BANK ANGLE MODE 1 PULL UP MODE 2 PULL UP MODE 2 PULL UP V1 ALERT ENGINE FAIL ALERT TERRAIN AWARENESS PREFACE TERRAIN AWARENESS WARNING OBSTACLE AWARENESS PREFACE OBSTACLE AWARENESS WARNING PWS Windshear Warning MODE 2 TERRAIN MODE 6 MINIMUMS MODE 6 ALTITUDE TERRAIN AWARENESS CAUTION OBSTACLE AWARENESS CAUTION MODE 4 TOO LOW TERRAIN TCF TOO LOW TERRAIN ALTITUDE ALERT MODE 6 ALTITUDE CALLOUTS SPEED BRAKE ALERT MODE 4 TOO LOW GEAR MODE 4 TOO LOW FLAPS MODE 1 SINKRATE MODE 3 DON‘T SINK MODE 5 GLIDESLOPE PWS Windshear Caution MODE 6 APPROACHING DH MODE 6 BANK ANGLE MODE 7 WINDSHEAR ALERT AUTOPILOT ALERT TAKEOFF FLAP ALERT TCAS RA TCAS TA ( RWS ) ( PWS ) (SIREN) WINDSHEAR WINDSHEAR WINDSHEAR BANK ANGLE PULL UP PREFACE TERRAIN TERRAIN PULL UP V1 ENGINE FAIL TERRAIN TERRAIN PULL UP OBSTACLE OBSTACLE PULL UP 1, 3 N/A − PWS Voice TERRAIN SELECTED CALLOUT ALTITUDE ALTITUDE CAUTION TERRAIN (PAUSE) CAUTION TERRAIN (7 sec pause) CAUTION OBSTACLE (PAUSE) CAUTION OBSTACLE (7 sec pause) TOO LOW TERRAIN TOO LOW TERRAIN 960 HZ TONE SELECTED CALLOUTS SPEED BRAKE (PAUSE) SPEED BRAKE TOO LOW GEAR TOO LOW FLAPS SINKRATE DON‘T SINK (PAUSE) DON‘T SINK GLIDESLOPE N/A − PWS Voice SELECTED CALLOUT BANK ANGLE (PAUSE) BANK ANGLE (QUIET) (or CAUTION WINDSHEAR if Caution voice Enabled) AUTOPILOT FLAPS (PAUSE) FLAPS N/A − TCAS Voice (may coincide with Mode 6 voices) N/A − TCAS Voice (may coincide with Mode 6 voices) Figure 91 HAM US/F-4 SaR Dec 2005 Basic Audio Menu Page 181 Lufthansa Technical Training M13.4 NAVIGATION EGPWS Part-66 MKV EGPWS PROGRAMMING Accomplishment Instructions General 1. Airplane must be electrically energized. 2. The PCMCIA card AIIiedSignaI with the actual version of Terrain database must be used. For Training Purposes Only Programming Perform Ioading procedure according to AIIied SignaI . 1. Get access to the EGPWC 2. Ensure that the 115VAC oircuit breaker to the EGPWC is ON, and that the COMPUTER OK LED on the EGPWC front panel is ON. 3. Open the door on the EGPWC front panel. 4. Insert the PCMCIA card into the PCMCIA card slot. Insert card according to the picture on the front panel of the EGPWC. 5. WhiIe the Ioading is in progress, the IN PROG LED remains ON and the COMPUTER OK LED is OFF. 6. When the Ioading is complete, the XFER COMP LED goes ON. 7. Remove the PCMCIA card from the EGPWC. 8. After approx. 1 5 seconds, the COMPUTER OK LED goes ON to indicate that the contents of the PCMCIA card were successfully Ioaded. 9. Close the door on the EGPWC front panel. HAM US/F-4 SaR Dec 2005 Page 182 Lufthansa Technical Training M13.4 NAVIGATION EGPWS Part-66 Top of card when inserted PCMCIA card slot For Training Purposes Only Bottom of card when inserted EGPW COMPUTER Figure 92 HAM US/F-4 SaR Dec 2005 EGPWC Page 183 Part-66 EGPWS: SELF TEST In addition to power−up and continuous BITE, user activated tests, via discrete test switches, and/or maintenance system commands are supported. Cockpit Self Test In aircraft with a cockpit−test switch it is possible to manually initiate tests and BITE annunciation while the aircraft is on ground. If the aircraft is above 2000 feet AGL cockpit self−test can be initiated, provided no warning or alert is in progress. A test switch on the unit‘s front panel, along with an audio headset jack, is also provided to give the flexibility of running tests both in the cockpit and at the LRU. For some maintenance systems the test command is input via an ARINC 429 input word. Six levels of information are available through voice messages by pressing the self−test switch. The test sequence in general is summarized as follows. S Level 1, Go/No−Go Test: This sequence indicates the systems ability to perform all of its configured functions. − Training Information Point You can start a level one self test from the front panel of the GPWC, but you cannot see the flight deck annunciations. Use the GPW module to start an operational test of the GPWS. − Short Level One Test — Normal Indications For this sequence, when the test switch is activated, the cockpit lamps are activated and aural messages are issued to indicate what functions are correctly operating. For instance, if no faults exist on an installation that uses the Terrain Awareness function in addition to basic GPWS and Windshear, then the result of the present status test would typically be: S ND system message TERR TEST shows in cyan S Ground proximity warning light on for 7 seconds S Glideslope lights on for 7 sec and aural message S Pull—up message displayed for 7 sec and aural message S Windshear message displayed for 7 sec and aural message S Terrain test pattern displayed on the ND and aural message S aural message sequence: Glideslope−−−−Pull Up−−−−Windshear Windshear Windshear−−−−Terrain Terrain, Pull Up However, if no valid Glideslope input were present, then the sequence would be S Glideslope INOP−−−−−Pull Up−−−−−Windshear Windshear Windshear−−−−Terrain Terrain, Pull Up During system self−test all INOP type visual annunciation are activated. − Long Level 1 Test For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS To start the long level 1 test, push and hold the test button until the first voice is heard. This will do the short level one test procedures then continue to give voice call−outs for all customer selected items. HAM US/F-4 SaR Dec 2005 Page 184 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 93 HAM US/F-4 SaR Dec 2005 EGPWS: SELF TEST, LEVEL 1 Page 185 Part-66 Self-test Level 2−6 Self−test levels 2−6 access is through the GPW module and the GPWC. When you use the GPWC, a 600 ohm headphone is necessary to listen to the test information. Plug the headphone into the jack on the front panel of the GPWC. If you do the tests from the flight deck, the information is heard over the flight deck speakers. Use the self−test button on the front panel of the GPWC or the self−test button on the GPW module to get access to levels 2−6. The self−test buttons have these two modes: − Short cancel — push the button for less than two seconds − Long cancel — push the button for more than two seconds. Use the self−test buttons for these functions: − Start self−test level 1 − Go to the next item or flight leg within a test − Go to the next self−test level − End the self−test. When a test level ends, the aural message PRESS TO CONTINUE annunciates. Push the self−test button to go to the next test level. If you do not push the self− test button within three seconds, self−test ends. S Level 2, Current Faults: This sequence identifies all faults, if any, that currently exist. It will distinguish between internal and external faults. If no faults exist, the message is „No Faults„. S Level 3, Configuration Information: This sequence indicates the versions of the resident hardware, software and database versions. Also provided is the current program pin option selections, including voice and callouts menu selections. S Level 4, Fault History: This sequence indicates all system faults that were logged for the past ten flight legs. (Information on the last 64 legs is accessible via the RS−232 interface). S Level 5, Warning History: This sequence provides all EGPWS alerts/warnings that were logged for the past ten flight legs. (Information on the last 64 legs is accessible via the RS−232 interface). For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 186 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 94 HAM US/F-4 SaR Dec 2005 EGPWS: SELF TEST Page 187 Part-66 S Level 6, Discrete Test: This sequence provides discrete input transitions as a to aid system installation and maintenance. The test starts with the aural message DISCRETE TEST. If the condition of a discrete input changes, you hear the new state of the discrete. You hear the aural message DISCRETE INPUT TEST — PRESS TO CANCEL every 60 seconds. Push a short or long cancel to end the self−test. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 188 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 95 HAM US/F-4 SaR Dec 2005 EGPWS: SELF TEST Page 189 Part-66 EGPWS: STATUS LEDS Purpose There are three status LEDs on the front panel of the ground proximity warning computer (GPWC). These LEDs turn on when there is power to the GPWC. These are the LEDs: − External fault — yellow − Computer OK — green − Computer fail — red. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS HAM US/F-4 SaR Dec 2005 Page 190 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION EGPWS Figure 96 HAM US/F-4 SaR Dec 2005 EGPWS: COMPUTER STATUS LEDs Page 191 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 Appendix A: Definitions The following acronyms are used in this document: Acronym Interpretation AAAS Alternate Audio Alert Select ADC Air Data Computer ADS Air Data System AGL Above Ground Level AHRS Attitude Heading Reference System AIMS Airplane Information Management System AOA Angle of Attack ASL Above Sea Level ATP Acceptance Test Procedure BCD Binary Coded Decimal BIST Built in Self Test BIT Built In Test BITE Built In Test Equipment BNR Binary BOSS Batch Oriented Simulation System C/O Callouts CAA Civil Aviation Authority CAIMS Central Aircraft Information/Maintenance System CFDS Centralized Fault Display System CFIT Controlled Flight Into Terrain CFM Cubic Feet per Minute CISRD CFDS Interface System Requirements Document CMC Central Maintenance Computer COTS Commercial Off the Shelf CP Control Panel CRS Course CW Clockwise DAA Digital/Analog Adapter DADC Digital Air Data Computer DAU Data Acquisition Unit DC Digital Command DDM Difference in Depth of Modulation DEVN Deviation DH Decision Height DITS Digital Information Transfer System DME Distance Measuring Equipment DO Discrete Output HAM US/F-4 SaR Dec 2005 DSP DSU DSWC EEPROM EFCP EFIS EGPWC EGPWD EGPWS EICAS EMI ENB EPROM F/T FAA FCC FDR FIAS FMC FMS FPM FSEU FWC G/S GMT GPS GPW GPWS GT H/W HDG HDOP HSID I/O IAC ICD ILS INOP IOC IRS Digital Signal Processor Display Switching Unit Digital Stall Warning Computer Electrically Erasable Programmable Read Only Memory EFIS Control Panel Electronic Flight Instrument System Enhanced Ground Proximity Warning Computer Enhanced Ground Proximity Warning Display Enhanced Ground Proximity Warning System Engine Indication and Crew Alert System Electromagnetic Interference Enabled Erasable Programmable Read Only Memory Functional TestF/W Fail/Warning Federal Aviation Administration Flight Control Computer Flight Data Recorder Flight Inspection Aircraft System Flight Management Computer Flight Management System Feet per Minute Flaps/Slats Electronic Unit Fault Warning Computer Glideslope Greenwich Mean Time Global Position System Ground Proximity Warning Ground Proximity Warning System Greater Than Hardware Heading Horizontal Dilution of Position Hardware/Software Interface Document Input/Output Integrated Avionics Computer Interface Control Document Instrument Landing System Inoperative Input/Output Concentrator Inertial Reference System Page 192 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION EGPWS Part-66 ISO IVS KT KTS LED LRRA LRU LSB LT MCP MDA MFD MKII MKVII MKV MLS MSB N/A NCD ND NVM OMS P/N PAR PC PCMCIA PFD PMAT PP PPI PVM QFE QNH RA RAM RDOP ROM RTCA RTS RWY International Standards Organization Inertial Vertical Speed Knots Knots Light Emitting Diode Low Range Radio Altimeter Line Replaceable Unit Least Significant Bit Less Than Mode Control Panel Minimum Barometric Altitude Multi−Functional Display Mark Two Warning Computer Mark Seven Warning Computer Mark V Warning Computer Microwave Landing System Most Significant Bit Not Applicable No Computed Data Navigation Display Non Volatile Memory Onboard Maintenance System Part Number Parity Personal Computer Personal Computer Memory Card Industry Association Primary Flight Display Portable Maintanence Access Terminal Program Pin Plan Position Indicator Processor/Voice/MemoryPWS Predictive Windshear System Corrected Baro Alt relative to field elevation Corrected Baro Alt relative to sea level Radio Altitude Random Access Memory Radar Display Output Processing Read Only Memory Requirements and Technical Concepts for Aviation Ready to Transmit Signal Runway HAM US/F-4 SaR Dec 2005 S/T S/W SDI SDRD SFFS SIG SPC SRD SSM ST SWC TACAN TAD TA&D TBD TCAS TCF TK TLB TSO TTL UART USM UTC UUT VDC VDOP VHF VLSI VOR W/S WC Self Test Software Source/Destination Identifier Software Design Requirements Document System Flight Fault Summary Significant Stall Protection Computer System Requirements Document Sign Status Matrix Self Test Stall Warning Computer Tactical Air Navigation Terrain Awareness Display Terrain Awareness & Display To Be Determined Traffic Collision Avoidance System Terrain Clearance Floor Track Translation Lookaside Buffer Technical Standing Order Tuned To Localizer Universal Asynchronous Receiver Transmitter Unsigned Magnitude Universal Time Correlation Unit Under Test Volts, DC Vertical Dilution of Precision Very High Frequency Very Large Scale Integrated Circuit VHF Omni−directional Range Windshear Warning Computer Page 193 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 RADAR FUNDAMENTALS GENERAL The word RADAR is an acronym formed from the words: RAdio - Detection - And - Ranging. Radar is a means of employing radio waves to detect and locate material objects such as aircraft, clouds or land masses. Location of an object is accomplished by by determining the distance and direction from the radar equipment to the object. The process of locating objects requires, in general, the measurement of three coordinates: S range S angle of azimuth and S angle of elevation. The characteristics of the object ( target ) can be of interest and are measured by processing the S amount of reflected energy, S amount and/or spectrum of dopplershift. In general the problems or duties can be solved with any frequency or wavelength, but for many reasons the most common wavelengths are limited between 1m and some mm. The most common radar- bands are assigned with letters: Band Frequency - f Wavelength - P 225 - 390 MHz 133.3 - 76.9 cm L 390 - 1550 MHz 76.9 - 19.3 cm S 1.55 - 5.2 GHz 19.3 - 5.7 cm C 5.3 - 5.8 GHz 5.7 - 5.2 cm X 5.2 - 10.9 GHz 5.8 - 2.8 cm Ku 10.9 - 36 GHz 2.6 - 0.8 cm Q 36 - 46 GHz 8.3 - 6.3 mm HAM US/F-4 SaR Dec 2005 Frequency The choice of frequency is determined by the purpose of the radar and the different characteristics of the frequency bands. The higher the frequency (smaller the wavelength) the larger is the backscatter cross−section per unit volume of the target hence the greater the echo power. However, high frequencies suffer more atmospheric absorption than do low, and further cannot penetrate clouds to the same extent. Thus the choice of frequency is a compromise. An additional consideration is the beamwidth; for a given antenna diameter a narrower beam is produced with a higher frequency. The choice of frequency for Weather Radar ( WR ) Systems is e.g. either about 3.2 cm (X−band) or 5.5 cm (C−band). The majority of WR Systems in service are X−band. Radar Classification Radar systems can be classified into S primary- secondary- radar and S pulse- continuous wave- radar. S Primary Radar System ( PRS ) The target only reflects a portion of the radarsignal- ( passive reflector )-. S Secondary Radar System ( SRS ) The transmitted radarsignal activates a transponder in the target. The transponder can modulate the signal and sends it back to the interrogator- ( active reflector )-. S Pulse Radar System The characteristics of a pulse system are the short transmission pulses. After transmission, the system waits for the echoes. This process is done many times a second and is called the Pulse Repetition Time ( PRT ). S Continuous Wave ( CW )Radar System In opposite to the pulse system, the CW system is continuously transmitting and receiving. Page 194 Part-66 THIS PAGE INTENTIONALLY LEFT BLANK For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 195 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 RANGE DETERMINATION The distance ( range ) to the target is determined by the time for the signal to travel to the target and return. If a signal is produced ( e.g. sound, radio wave or light ) which travels at a velocity of v in direction of a reflecting object, the echo from this target will return at the transceiver after the traveltime TL which the signal needs for the twoway distance between transceiver and target. If the velocity v is known, the traveltime TL is a value for the distance R. + ; + ; Radarequipments transmit radiosignals. The velocity v of this electromagnetic waves is about 300.000 km/sec or 162.000 nm/sec. The traveltime TL for an object, that is 1 nautical mile away from the transceiver is 12,36 s. This TL of 12,36 s very often is called a radar mile. 1 Radar Mile ¢ 12,36 s If the distance between radar and target is 1 km, the traveltime is 6,66 s. Modulation of the radarsignal To be able to measure the traveltime TL the transmit-signal must be modulated, so that the correlation between transmit- and receive- signal can be achieved. S Pulse Radar The characteristic of a pulse radar is, that it sends only short pulses of some microseconds, which are seperated by longer pauses of some milliseconds. Within this pauses, the system is listening for echoes. At elder systems, the pauses - and they are expressed by the Pulse Repetition Time ( PRT ) , must be long enough to assign transmit- and receive- signal. At the latest radar generations, the correlation is done by a JITTER- or DITHER- Technique. S CW Radar To achieve the correlation between transmit- and receive- signal by using a CW Radar, it is necessary that: a. there is a relative movement between radar and object or b. the transmit- frequency must be changed continuously ( FM ). In case a. the receive- signal contains the dopplershift, so that transmit- and receive- signal can be assigned. Because the dopplershift is independent from the distance, a range measurement is not possible. For Training Purposes Only In case b. the transmit- signal is frequency modulated with a triangle- or sawtooth- signal ( FM-CW- Radar ). Due to the modulation, there will be a frequency difference Df between transmit- and receive- signal. The amount of frequency difference is a function of the traveltime TL and therefore of the range. HAM US/F-4 SaR Dec 2005 Page 196 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Pulse Radar PRT t XMIT Echo target 1 Target XMIT target 2 target 3 receive v transmit transmit CW Radar RCV F For Training Purposes Only R Distance between transmitter and target Df TL t Figure 97 HAM US/F-4 SaR Dec 2005 Radar Range Determination Page 197 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 RADAR SYSTEM PARAMETERS The signal propagation in the used cm-range, is given by the optical law of the “line of sight“. Due to effects within the atmosphere, the distance for radiowaves is more than for lights, so that the radar horizon is behind the optical horizon. Beside the limitation of the radar horizon, the maximum range of radar systems is mainly determined by the following parameters: S transmit power S echo power S attenuation and reflectivity S receiver sensitivity S pulse repetition Transmit Power To receive a sufficient echo signal from a small target over the maximum distance, the transmit power must be high. The radar transmitter generates RF energy in the form of short pulses or as continuous signals. The pulse system is able to produce a high peak power Pi . The relationship between the average power and the peak power is a function of pulse width ti and pulse repetition time PRT. For Training Purposes Only ǒ )%. "*%Ǔ ǒ(&. '/'/Ǔ + ǒ. "*%Ǔ ǒ(&. #''"!. . Ǔ The relation between pulse width and pulse repetition time PRT is called the Duty Cycle. ǒ )%. "*%Ǔ (',. ,/ +/ ǒ. "*%Ǔ If the power is transmitted via a directional antenna, the power is concentrated in a narrow beam, and the result will be a much higher transmit power in direction of the target due to the gain of the antenna. The antenna gain for radar antennas is referenced to an isotropic radiator. + . "/ ǒ%%. !'!!Ǔ . ǒ &"'%"#. %'"%Ǔ The antenna gain for a parabolic antenna e.g. ANT-1T is minimum 33 dB. The antenna gain for a flat plate antenna e.g. DAA-4A is minimum 35 dB. Echo power The echo power is - for a given transmit power - a function of S the distance to the target and S the radar cross section Because the power is transmitted into a sphere, the power intensity at the target is given by : / +/ p The reflected power P target depends on the characteristics of the target. The effective cross section for the mostly complex targets is called the radar cross section. The reflected power from the target is again distributed onto the surface of a sphere, so that the received echo power intensity again is a function of 4 R 2. As a result, the echo power decreases with 1/ R 4. A typical duty cycle for magnetron transmitter is 1 : 1000. The reciprocal of the PRT is the pulse repetition frequency PRF. / +/ HAM US/F-4 SaR Dec 2005 Page 198 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 RF power ÉÉÉÉ ÉÉÉÉ ÉÉÉÉ ÉÉÉÉ ÉÉÉÉ ÂÂÂÂ ÇÇÇÇÇÇÇÇÇÇ ÉÉÉÉ ÂÂÂÂ ÇÇÇÇÇÇÇÇÇÇ ÉÉÉÉ ÉÉÉÉ ÂÂÂÂ ÇÇÇÇÇÇÇÇÇÇ ÂÂÂÂ ÇÇÇÇÇÇÇÇÇÇ sphere surface 4 R 2 peak power equal areas power t resting time pulse repetition time For Training Purposes Only pulse width ANT average Figure 98 HAM US/F-4 SaR Dec 2005 Radar Transmit- / Echo- Power Page 199 Part-66 Attenuation and Reflectivity Till now we were locking only to an ideally propagation area, i.e. without any attenuation. Practically the radar waves are attenuated by penetrating air or moisture. This attenuation is a function of the frequency and increases with higher frequencies. Rain, clouds, snow or hail increase the attenuation and therefor lower the distance. If the radar shall be used to lock thru the clouds, the frequency is not allowed to be too high, otherwise the attenuation will be too high. On the other hand, if the system shall be used to show the intensity of rainfall within a cloud and also the dimenson of the cloud, there has to be a compromise between high reflectivity (high frequency) and low penetration attenuation (low frequency). The compromise for WXR systems was found at X- Band ( X 9GHz ). For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 200 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 REFLECTIVITY HEAVY CLOUD BURST HEAVY RAIN LIGHT RAIN For Training Purposes Only DRIZZLE AMOUNT OF RAIN REFLECTIVITY IN RAIN CLOUDS Figure 99 HAM US/F-4 SaR Dec 2005 ATTENUATION IN RAIN CLOUDS Radar Attenuation and Reflectivity of E.M.Waves Page 201 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Receiver Sensitivity The radardistance is limited by the receiver sensitivity. The limit is given at the distance, where the received echo power Pe and the noise power of the receiver are of the same value. Pe = K x T0 x B K = Bolzmann-Constant 1,37 x 10 -23 ƪWs/Kƫ T0= Noisetemperature of the receiver ƪKƫ B = Receiver bandwidth ƪHzƫ In the pulse radar technology, the connection between bandwidth B and pulseduration ti is given by : B x ti [ 1 If the pulsetime ti is decreased e.g. for a better solution, the bandwidth B must be increased; this will reduce the receiver sensitivity. For a high receiver sensitivity, the pulsetime should be as high as possible. The receiver sensitivity for radar systems is given by the term Minimum Discernable Signal ( MDS ) - Level. The MDS-Level for the Bendix RDR-1F is -104 dBm. The MDS-Level for the Bendix RTA-4A is -122 dBm. Pulse Repetition The Pulse Repetition Frequency ( PRF ) or Pulse Repetition Time ( PRT ) effects the radardistance in the following respects: 1. To maintain a constant duty cycle ( pulse width x PRF ), an increase in PRF must be accompanied by a decrease in pulse width. 2. The minimum necessary echo power will be lower, if the target will be hidden as often as possible. The number of hits will increase, the higher the PRF will be for a given antenna beamwidth and antenna sweepfrequency. 3. Because the echoes from the maximum distance has to be received before the next interrogation is started, the selected PRF must guarantee, that for the wanted range only one pulse is „ in the air „. The maximum PRF is calculated from the maximum pulse travel time. 0 +0 + If the max range shall be e.g. 4. 400 nM, the maximum PRF is : & +& ƪ mń = 200Hz For Training Purposes Only Between the contrary demands of 2. and 3. there has to be found a compromise. In the latest radar generation - like FLW - this problem is solved by using variable PRFs. The PRF is changed by JITTER - or DITHER - technology . This technology makes it possible, to get the necessary high number of hits for e.g. windshear detection and solve the problem of multible trace echoes. HAM US/F-4 SaR Dec 2005 Page 202 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Pulse Radar PRTmin PRTwrong XMIT ÉÉ ÉÉ ÉÉ ÉÉ t Echo target 1 target 2 target 3 receive For Training Purposes Only transmit Figure 100 HAM US/F-4 SaR Dec 2005 Radar PRF Page 203 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 RADAR SOLUTION The solution of a pulse radar is determined by: S pulse width ti S beam width S antenna sweep frequency Pulse Width ti The pulse width limits the: S radial accuracy S range solution S minimum range The use of long pulses, will result in high receiver sensitivities and thus improved range. There are arguments against long pulses: 1. Radial accuracy deteriorates with increasing pulse width. A pulse of 1 s will produce an echo of 1 s from a point target. Because 1 s corresponds to a distance of 150 m, the point target will be displayed as a 150 m target. For a high radial accuracy, the pulse width must be as short as possible. 2. Range solution deteriorates with increasing pulse width. A pulse of 1s duration occupies about 300 m in space. If two targets are on the same bearing but within 150 m one behind the other, the echo from the nearest target ( target echo #1 ) is still being received when the leading edge of the echo from the next target ( target echo #2 ) is received. The result is that both targets merge on the display. The range of the targets does not affect the solution. For a high range solution, the pulse width must be as short as possible. 3. Since there is only one antenna and a common frequency for transmitter and receiver, the antenna must be switched to the transmitter for the duration of the pulse; thus the pulse width determines minimum range. For a 2s pulse no return can appear for the first 2s of the time−base, so the minimum range will be : c x ti / 2 300 m. For a short minimum range, the pulse width must be as short as possible. Because the above mentioned arguments for a short pulse width are more important for selected short ranges than for long ranges, the modern radar system changes the pulse width ( and also the band width of the receiver ) with the selected range. HAM US/F-4 SaR Dec 2005 Page 204 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 target antenna time echo transmit pulse point targets For Training Purposes Only target echo Figure 101 HAM US/F-4 SaR Dec 2005 target echo Radar Solution Criterion: Pulse Width Page 205 Part-66 Beam Width A narrow beam is always preferred, because it will increase the antenna gain. But a narrow beam needs a large antenna diameter ( limited by the radom ) or a high frequency ( limited by the attenuation- and reflectivity- aspects ). Also the minimum number of hits on a target to receive a useful echo, limits the beam width ( at a given PRF and antenna sweep frequency ). The beam width limits: S accuracy in azimuth S solution in azimuth 1. Accuracy in the azimuth deteriorates with increasing beam width and distance. For a beam width of 3 a point target at a S distance of 20 km will be shown as 1 km and at a S distance of 200 km shown as 10 km wide target. The accuracy in azimuth is - contrary to the radial accuracy - strongly dependent from the distance. 2. Simple geometric considerations show, that with a 3 beam width, two targets separated by about 2,5nautical miles, at a range of 50 nautical miles, will appear as one on the indicator. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 206 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 target 1 Two targets displayed as one TARGET target 1 one target displayed wider in azimuth due to : - beam width and - distance For Training Purposes Only target 2 target 2 Two targets displayed separately Figure 102 HAM US/F-4 SaR Dec 2005 Radar Solution Criterion: Beam Width Page 207 Part-66 WAVE GUIDES General To transfer electromagnetic waves, four different possible methods are in use: 1. single wire over ground ( asymmetrical line ) 2. parallel lines ( symmetrical line ) 3. coaxial cable 4. wave guide The single wire over ground has losses due to S ohmic resistance S skin effect S radiation loss S dielectric losses The same applies for the parallel lines but with lower radiation loss. Coaxial cables have no radiation losses due to the shielding but at higher frequencies the skin effect becomes more effect. The wave guide is the optimum solution for radio wave transmission at high frequency, because it has only little ohmic resistance, skin effect, radiation- and dielectric- losses . Unfortunately, the necessary dimensions are depending on the transferred wave length. The wide dimension of the wave guide must be greater than /2. So wave guides are only in use for microwaves. The narrow dimension determines the power capability. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 208 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 shield dielectric displacement currents conductor currents conductor For Training Purposes Only Momentary Picture of a Lecher Wave in a Coaxial Cable Momentary Picture of a Wave in a Waveguide Figure 103 HAM US/F-4 SaR Dec 2005 Radar Coaxial Cable and Wave Guide Page 209 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Modes of operation Generally there can be an infinite number of different wave forms stimulated in a wave guide. Common to all of them is, that only one component of the electro-magnetic field, either the E- or the H- component is in the direction of the axis. The main characteristic used for identification of the wave form, S in Europe is the longitudinal field component and S in US the lateral field component. If only the H-Field is in direction of the travel axis, S this is designated as H- or TE- mode. If only the E-Field is in direction of the travel axis, S this is designated as E- or TM- mode. In addition to the above designation, subscript numbers are used to complete the discription of the field pattern. rectangular wave guide circular wave guide 1. Index number of /2 patterns in the wide dimension number of patterns around the circumference 2. Index number of /2 patterns in the narrow dimension number of /2 patterns along the diameter For Training Purposes Only Except for special uses, the common wave guide is the rectangular one. The mode H10 or TE 10 is the one with the longest wavelength and therefore it is the most common, because it has the lowest attenuation. HAM US/F-4 SaR Dec 2005 Page 210 Part-66 H 10 / TE 10 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS H 20 / TE 20 H 11 / TE 11 E 01 / TM 01 Figure 104 HAM US/F-4 SaR Dec 2005 H 11 / TE 11 H 01 / TE 01 Radar Modes in Wave guides Page 211 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Wave propagation in wave guides The travel of energy in a wave guide is similar to but not identical to that of the propagation in free space. If energy is to be propagated through a wave guide, two boundary conditions must be met: S no electrical field components are allowed to be parallel to the surface of the wave guide S no magnetic field components are allowed to be perpendicular to the surface of the wave guide To meet the above conditions, the wave front within a wave guide can not travel through it straight forward. When a small probe is inserted in a wave guide, and excited with RF energy, it will act as a vertical antenna. Positive and negative wave fronts will be radiated from the probe, as shown in Figure 105 . The portion of any wave front traveling in direction B will be rapidly attenuated because it will not fulfill the required boundary conditions. The portions of any wave front traveling in directions A or C will resolve themselves into oblique wave fronts traveling across the guide. This action is also illustrated in Figure 105 . Note that, wave fronts of the same polarity cross at the center of the guide. This produces a maximum E−field along the center of the guide. Also, opposite polarity wave fronts meet at the walls of the guide, thereby cancelling. This causes the E− field to be zero at the walls of the guide. The foregoing action fulfills the boundary conditions previously stated. The angle at the point at which the wave front strikes the wall of the guide and the perpendicular to the surface at the point of arrival is known as the ANGLE OF INCIDENCE ( ). The angle at which the wave front is reflected from the wall of the guide and the normal is called the ANGLE OF REFLECTION ( ) . These angles are equal, and are depicted in Figure 105 . Figure 105 illustrates the angles of incidence and reflection for a high frequency and a low frequency just above the cutoff frequency f0. HAM US/F-4 SaR Dec 2005 The angles of incidence and reflection are a function of frequency (assuming the guide dimensions are held constant). For the angles and the following equations are given: # !a + ǒ lń Ǔ = l ; "# b + ǒ lń Ǔ = l ; As the frequency decreases, the angles increase. is equal to 2a, sin and cos will be For this frequency -called the cutoff frequency f0- angle is 90 is 0. For the frequency, at which the Therefore, at any frequency lower than f0, the wave fronts will be reflected back and forth across the guide, setting up standing waves and no energy will be conducted down the guide. Page 212 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 A B ANTENNA C valley mountain For Training Purposes Only high frequency low frequency Figure 105 HAM US/F-4 SaR Dec 2005 Radar Wave Propagation in Wave Guides Page 213 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Velocities in wave guides In Figure 106 , although the wave front is traveling with velocity of light Vc, it is not moving straight down the guide. Its straight line velocity or axial velocity appears to be less than the speed of light. This axial or straight line velocity is called the Group Velocity (VG). The relationship of the group velocity to the diagonal velocity illustrates an unusual phenomenon. Referring to Figure 106 , during a given time, a wave front will move from point A to point B, at the Velocity of light (Vc). Due to this diagonal movement, the wave front actually has moved down the guide only a shorter distance. The velocity with which the wavefront has moved through this distance is the group velocity (VG). However, if an instrument were used to detect the two positions 1 and 2 of the phase at the wall, the distance is greater than distance for VC or VG. Therefore, the contact point between the wave front and the wall appears to be moving with a velocity greater than the velocity of light. Since the phase of the RF has changed over distance , this velocity is called the Phase Velocity (V ). The mathematical relationship between the three velocities is given by the equation: For Training Purposes Only + Ǹǒ fǓ "# a +% ; + "# a "# a +% ; f f + "# a Group velocity decreases with a decrease in frequency, and phase velocity increases with a decrease in frequency. For the cutoff frequency, is 90 and therefore cos is 0 That means, that at the cutoff frequency f0 : S the group velocity will be VG = 0 and S the phase velocity will be V = R HAM US/F-4 SaR Dec 2005 Page 214 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ B ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ A 2 1 For Training Purposes Only VG V Figure 106 HAM US/F-4 SaR Dec 2005 Radar Velocities in Wave Guides Page 215 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Wave guide junctions It is practically impossible to construct a waveguide system in one piece. It is necessary to construct guide sections, which must be connected by joints. Any irregularities in the joints cause standing waves, and permit energy to be dissipated. A good connection between the wave guide parts is especially needed at those sections, where currentflow at the wave guide surface is high. Figure 107 shows the wall current distribution for the most common H10 mode in a rectangular wave guide. A proper permanent joint affords a good connection between the parts of a waveguide and has very little effect on the fields. Normally this type of joint is made at the factory. When it is used, the waveguide sections are machined within a few thousandths of an inch and then welded together. The result is a hermetically sealed and mirror smooth joint. Choke joint Where sections of waveguide must be taken apart for normal maintenance and repair, it is impractical to use a permanent joint. To permit portions of the waveguide to be separated, the sections are connected by semipermanent joints of which the choke joint is the most common type. A crosssection view of a choke joint is shown inFigure 107 . It consists of two flanges which are connected to the waveguide at the center. The top flange is flat, and the bottom one is slotted one−quarter wavelength deep from the inner surface of the waveguide, at a distance of one−quarter wavelength from the point where the flanges are joined. The quarter−wavelength slot is short−circuited at the end. The two quarter−wavelength sections form a halfwavelength section and reflect a short−circuit at the place where the walls are joined together. Electrically, this creates a short−circuit at the junction of the two waveguides- especially at those areas, where the maximum wall currents are in direction of the wave guide. The two sections may actually be separated as much as a tenth of a wavelength without excessive loss of energy at the joint. This separation allows room to seal the interior of the waveguide with a rubber gasket for pressurization. The loss introduced by a well designed choke is less than 0.03 db, while an unsoldered permanent joint, well machined, has a loss of 0.05 db or more. HAM US/F-4 SaR Dec 2005 Rotating joint Rotating joints are usually required in a radar system where the transmitter is stationary and the antenna is rotatable. A simple method for rotating part of a waveguide system uses a mode of operation that is symmetrical about the axis. This requirement is met by using a circular wave guide and a mode such as E01/TM01. In this method, a choke joint may be used to separate the sections mechanically and to join them electrically. The waveguide rotates, but not the field. The fixed field minimizes reflections. This is basically the method used for all rotating joints. Since most radars use rectangular waveguides,the circular rotating joint must be inserted between two rectangular sections. The E−lines H10/TE10 couple from the rectangular guide into the circular guide and excite the circular guide in the E01/TM01 mode. This is the mode that provides the required axial symmetry for rotating joints. At the top of the joint, the E−lines couple the energy into the rectangular guide which leads to the antenna. Here the guide is operating again in the H10/TE10 mode. Another method is to use the part of a coaxial cable for rotation. The energy from the rectangular wave guide is picked up by an antenna formed by the inner conductor of a coax, brought to the rotational rectangular wave guide, and is again transmitted by the inner conductor of the coax into the wave guide. Page 216 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 rotational fix momentary picture of a H10 - wave wall currents of a H10 - wave flat flange wave guide A rotating circular wave guide coupler contact area wave guide B rotational choke flange For Training Purposes Only fix blocking flange Lecher wave in the coaxial cable rotating coaxial coupler wave guide junction Figure 107 HAM US/F-4 SaR Dec 2005 Radar Wave Guide Junctions Page 217 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Coupling methodes Fundamentally, there are three methods of coupling energy into or out of a waveguide: S PROBE-Coupling S LOOP-Coupling and S SLOT-Coupling. Probe- or capacitive- coupling is illustrated in Figure 108 A. Its action is the same as that of a quarter−wave antenna. When the probe is excited by an RF signal, an electric field is set up. The probe should be located in the center and a quarter−wavelength, or odd multiple of a quarter−wavelength, from the short−circuited end as illustrated. This is a point of maximum E−field and, therefore, is a point of maximum coupling between the probe and the field. Usually, the probe is fed with a short length of coaxial cable. The outer conductor is connected to the waveguide wall, and the probe extends into the guide, but is insulated from it.The degree of coupling may be varied by varying the length of the probe, removing it from the center of the E−field, or shielding it. In a pulse modulated radar system there are wide sidebands on either side of the carrier frequency. In order that a probe does not discriminate too sharply against frequencies which differ from the carrier frequency, a wide band probe may be used. For Training Purposes Only Loop- or inductive- coupling is illustrated in Figure 108 B. The loop is placed at a point of maximum H−field in the guide. As shown, the outer conductor is connected to the guide and the inner conductor forms a loop inside the guide. The current flow in the loop sets up a magnetic field in the guide. The loop may be placed in a number of locations. The degree of loop coupling may be varied by rotation of the loop. Slot- or aperture- coupling is illustrated in Figure 108 C. The third method of coupling is slot-, or aperture - coupling. Slots a are at an area of maximum wall-current rectangular to the slots and therefore they have a high degree of coupling. Slot b is at an area of minimum wall-current rectangular to the slot and therefore it has no coupling. Slot c is at an area of some wall-current rectangular to the slot and therefore it has some coupling. HAM US/F-4 SaR Dec 2005 Page 218 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 A : Probe- or capacitive- coupling B : Loop- or inductive- coupling C : Slot- or aperture- coupling For Training Purposes Only a a Figure 108 HAM US/F-4 SaR Dec 2005 a b c Radar Coupling methodes Page 219 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 MICROWAVE GENERATION Microwaves can be generated by: S Magnetrons S Klystrons S Reflexklystrons S Gunn Diodes S Impatt Diodes S Tunnel Diodes Klystrons and magnetrons are used in elder radar systems for high power generation, modern low power radars are semiconductor equipped and need no warming up. Description of the cavity magnetron The cavity (or traveling wave) magnetron high−power microwave oscillator is a diode which uses the interaction of magnetic and electric fields in a complex cavity to provide oscillations of very high peak power. The cavity magnetron is a diode, usually of cylindrical construction. It employs a radial electric field, an axial magnetic field and an anode structure with perm−anent cavities. The cylindrical cathode is surrounded by the anode with cavities, and thus a radial dc electric field will exist. The magnetic field, because of a magnet has lines of magnetic force passing through the cathode and the surrounding interaction space. The lines are thus at right angles to the structure cross section. The magnetic field is also dc, and since it is perpendicular to the plane of the radial electric field, the magnetron is called a crossed−field device. The output is taken from one of the cavities, by means of a coaxial line, or through a waveguide, depending on the power and frequency. The anode is normally made of copper, regardless of its actual shape. The magnetron has a number of resonant cavities and must therefore have a number of resonant frequencies and/or modes of oscillation. Whatever mode is used, it must be self−consistent. For example, it is not possible for the eight−cavity magnetron (which is often used in practice) to employ a mode in which the phase difference between the adjacent anode pieces is 30. If this were done, the total phase shift around the anode would be 8 x 30 = 240, which means that the first pole piece would be 120 out of phase with itself! Simple investigation shows that the smallest practical phase difference that can exist here between adjoining anode poles, is 45 or /4 rad, giving a self−consistent overall phase shift of 360 or 2 rad. HAM US/F-4 SaR Dec 2005 Effect of magnetic and electric fields When magnetic and electric fields act simultaneously upon the electron, its path can have any of a number of shapes dictated by the relative strengths of the mutually perpendicular electric and magnetic fields. Some of these electron paths are shown in the figure in the absence of oscillations in a magnetron, in which the electric field is constant and radial, and the axial magnetic field can have any number of values. When the magnetic field is zero, the electron goes straight from the anode to the cathode, accelerating all the time under the force of the radial electric field; this is indicated by path x. When the magnetic field has a small but definite strength, it will exert a lateral force on the electron, bending its path to the left. The electron’s motion is no longer rectilinear. As the electron approaches the anode, its velocity continues to increase radially as it is accelerating. Therefore, the effect of the magnetic field upon it increases also, so that the path curvature becomes sharper as the electron approaches the anode. It is possible to make the magnetic field so strong that electrons will not reach the anode at all. The magnetic field required to retum electrons to the cathode after they have just grazed the anode is called the cutoff field. The resulting path is z. Knowing the value of the required magnetic field strength is important because this cutoff field just reduces the anode current to zero in the absence of oscillations. If the magnetic field is stronger still, the electron paths as shown will be more curved still, and the electrons will retum to the cathode even sooner (only to be reemitted). All these paths are naturally changed by the presence of any RF field due to oscillations, but the state of affairs without the RF field must still be appreciated, for two reasons. First, it leads to the understanding of the oscillating magnetron. Second, it draws attention to the fact that unless a magnetron is oscillating, all the electrons will be returned to the cathode, which will overheat and ruin the tube. This happens because in practice the applied magnetic field is greatly in excess of the cutoff field. Oscillations are capable of starting in a device having high−Q cavity resonators. Page 220 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Electron paths without oscillations, with different magnetic field strengths For Training Purposes Only Cross section of a magnetron Bunched electron clouds rotating around the magnetrons cathode Paths traversed by electrons in a magnetron under -mode oscillations Figure 109 HAM US/F-4 SaR Dec 2005 Radar Magnetron Page 221 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Multicavity Klystron The two−cavity amplifier klystron forms a high−velocity electron beam, focused (external magnetic focusing is omitted for simplicity) and sent down a long glass tube to a collector electrode which is at a high positive potential with respect to the cathode. The beam passes gap A in the buncher cavity, to which the RF signal to be amplified is applied, and it is then allowed to drift freely, without any influence from RF fields, until it reaches gap B in the output or catcher cavity. If all goes well, oscillations will be excited in the second cavity which are of a power much higher than those in the buncher cavity, so that a large output can be taken. The beam is then collected by the collector electrode. The cavities are reentrant and are also tunable (although this is not shown). They may be integral or demountable. In the latter case, the wire grid meshes are connected to rings external to the glass envelope, and cavities may be attached to the rings. The drift space is quite long, and the transit time in it is put to use. However, the gaps must be so short that the voltage across them does not change significantly during the passage of a particular bunch of electrons; having a high collector voltage helps in this regard. It is apparent that the electron beam, which has a constant velocity as it approaches gap A, will be affected by the presence of an RF voltage across the gap. However, the extent of this effect on any one electron will depend on the voltage across the gap when the electron passes this gap. It is thus necessary to investigate the effect of the gap voltage upon individual electrons. Consider the situation when there is no voltage across the gap. Electrons passing it are unaffected and continue on the collector with the same constant velocities they had before approaching the gap . Sometime later, after an input has been fed to the buncher cavity, an electron will pass gap A at the time when the voltage across this gap is zero and going positive. Let this be the reference electron y. It is of course unaffected by the gap, and thus it is shown with the same slope on the Applegate diagram of the figure as electrons passing the gap before any signal was applied. Another electron, z, passes gap A slightly later than y, as shown. Had there been no gap voltage, both electrons would have continued past the gap with unchanged velocity, and therefore neither could have caught up with the other. Here, however, electron z is slightly accelerated by the now positive voltage across gap A, and given enough time, it will catch up with the reference electron. As shown on the Applegate diagram, it has enough time to catch electron y easily before gap B is approached. Similarly, electron x passes gap A slightly before the reference electron. However, although it passed gap A before electron y, it was retarded some- HAM US/F-4 SaR Dec 2005 what by the negative voltage then present across the gap. Since electron y was not so retarded, it has an excellent chance of catching electron x before gap B . As electrons pass the buncher gap, they are velocity−modulated by the RF voltage existing across this gap. Such velocity modulation would not be sufficient, in itself, to allow amplification by the klystron. However, electrons have the opportunity of catching up with other electrons in the drift space. When an electron catches up with another one, it may simply pass it and forge ahead. On the other hand, it may exchange energy with the slower electron, giving it some of its excess velocity, and the two bunch together and move on with the average velocity of the beam. As the beam progresses farther down the drift tube, so the bunching becomes more complete, as more and more of the faster electrons catch up with bunches ahead. Eventually, the current passes the catcher gap in quite pronounced bunches and therefore varies cyclically with time. This variation in current density is known as current modulation, and this is what enables the klystron to have significant gain. It will be noted from the Applegate diagram that bunching can occur once per cycle, centering on the reference electron. The limits of bunching are also shown. Electrons arriving slightly after the second limit clearly are not accelerated sufficiently to catch the reference electron, and the reference electron cannot catch any electron passing gap A just before the first limit. Bunches thus also arrive at the catcher gap once per cycle and deliver energy to this cavity. In ordinary vacuum tubes, a little RF power applied to the grid can cause large variations in the anode current, thus controlling large amounts of dc anode power. Similarly in the klystron, a little RF power applied to the buncher cavity results in large beam current pulses being applied to the catcher cavity, with a considerable power gain as the result. Needless to say, the catcher cavity is excited into oscillations at its resonant frequency (which is equal to the input frequency), and a large sinusoidal output can be obtained because of the flywheel effect of the output resonator. Page 222 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Klystron amplifier drift space for electron package generation For Training Purposes Only anode Applegate diagram for the klystron amplifier cathode accelerator grid input resonator output resonator Klystron Figure 110 HAM US/F-4 SaR Dec 2005 Radar Klystron Page 223 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Reflex Klystron It is possible to produce oscillations in a klystron device which has only one cavity, through which electrons pass twice. This is the reflex klystron. The reflex klystron is a low−power, low−efficiency microwave oscillator. It has an electron gun similar to that of the multicavity klystron but smaller. Because the device is short, the beam does not require focusing. Havingh been formed, the beam is accelerated towards the cavity, which has a high positive voltage applied to it and acts as the anode. The electrons overshoot the gap in this cavity and continue on to the next electrode, which they never reach. This repeller electrode has a fairly high negative voltage applied to it, and precautions are taken to ensure that it is not bombarded by the electrons. Accordingly, electrons in the beam reach some point in the repeller space and are then turned back, eventually to be dissipated in the anode cavity. If the voltages are properly adjusted, the returning electrons give more energy to the gap than they took from it on the outward journey, and continuing oscillations take place. As with the multicavity klystron, the operating mechanism is best understood by considering the behavior of individual electrons. This time, however, the reference electron is taken as one that passes the gap on its way to the repeller at the time when the gap voltage is zero and going negative. This electron is of course unaffected, overshoots the gap, and is ultimately retumed to it, having penetrated some distance into the repeller space. An electron passing the gap slightly earlier would have encountered a slightly positive voltage at the gap. The resulting acceleration would have propelled this electron slightly farther into the repeller space, and the electron would thus have taken a slightly longer time than the reference electron to retum to the gap. Similarly, an electron passing the gap a little after the reference electron will encounter a slightly negative voltage. The resulting retardation will shorten its stay in the repeller space. It is seen that, around the reference electron, earlier electrons take longer to retum to the gap than later electrons, and so the conditions are right for bunching to take place. In the multicavity klystron, velocity modulation is converted to current modulation in the repeller space, and one bunch is fanned per cycle of oscillations. It should be mentioned that bunching is not nearly as complete in this case, and so the reflex klystron is much less efficient than the multicavity klystron. HAM US/F-4 SaR Dec 2005 Transit time As usual with oscillators, it is assumed that oscillations are started by noise or switching transients. Accordingly, what must now be shown is that the operation of the reflex klystron is such as to maintain these oscillations. For oscillations to be maintained, the transit time in the repeller space, or the time taken for the reference electron from the instant it leaves the gap to the instant of its return, must have the correct value. This is determined by investigating the best possible time for electrons to leave the gap and the best possible time for them to return. The most suitable departure time is obviously centered on the reference electron, at the 180 point of the sine−wave voltage across the resonator gap. It is also interesting to note that, ideally, no energy at all goes into velocity−modulating the electron beam. It admittedly takes some energy to accelerate electrons, but just as much energy is gained from retarding electrons. Since just as many electrons are retarded as accelerated by the gap voltage, the total energy outlay is nil. This actually raises a most important point: energy is spent in accelerating bodies (electrons in this case), but energy is gained from retarding them. It is thus evident that the best possible time for electrons to retum to the gap is when the voltage then existing across the gap will apply maximum retardation to them. This is the time when the gap voltage is maximum positive. Electrons then fall through the maximum negative voltage between the gap grids, thus giving the maximum amount of energy to the gap. Therefore, the best time for electrons to retum to the gap is at the 90 point of the sine−wave gap voltage. Returning after 1.75 cycles obviously satisfies these requirements. Generally: T = n + 3/4 where T is the transit time of electrons in the repeller space, cycles, and n is any integer. Modes The transit time obviously depends on the repeller and anode voltages, so that both must be carefully adjusted and regulated. Once the cavity has been tuned to the correct frequency, both the anode and repeller voltages are adjusted to give the correct value of T from data supplied by the manufacturer. Each combination of acceptable anode−repeller voltages will provide conditions permitting oscillations for a particular value of n. In turn, each value of n is said to correspond to a different reflex klystron mode, practical transit times corresponding to the range from 1.75 to 6.75 cycles of gap voltage. Modes corresponding to n = 2 or n = 3 are the ones used most often in practical klystron oscillators. Page 224 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 tuning out cathode drift space reflector resonator For Training Purposes Only reflector voltage Figure 111 HAM US/F-4 SaR Dec 2005 Radar Reflex Klystron Page 225 Part-66 SEMICONDUCTOR MICROWAVE DEVICES To generate or amplify oscillations in microwave regions, magnetrons or clystrons are more and more substituted by semiconductor devices. The prerequisites for generation or amplification is to decrease the losses in resonant circuits represented by an ohmic resistor. To decrease the ohmic resistor - and by that also the losses - a device with a „ negative resistor „ is necessary. Three devices that have this characteristic are : S Gunn - Diode S Impatt - Diode S Tunnel - Diode All of the devices operate totally different, but they all have a negative resistance under various conditions. A negative resistor can be found e.g. S if the mobility of the electrons will decrease with increasing voltage (Gunn) S if a phasedifference of 180 between current and voltage exists (Impatt) S if the current decreases with increasing voltage (Tunnel) For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 226 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Gunn - Diode N) N N)) P) N C A Tunnel - Diode P) N) C A N) C For Training Purposes Only A Impatt - Diode Figure 112 HAM US/F-4 SaR Dec 2005 Microwave Semiconductors Page 227 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 Gunn Diode In 1963, Gunn discovered the ”transferred electron effect”. This effect is used in the generation of microwave oscillations in bulk semiconductor materials. The effect was found by Gunn to be exhibited by gallium arsenide ( and also some other materials ). If a relatively small dc voltage is placed across a thin slice of gallium arsenide, then negative resistance will manifest itself under certain conditions. Basically, these consist merely of ensuring that the voltage gradient across the slice is in excess of about 3300 V/cm. Oscillations will then occur if the slice is connected to a suitably tuned circuit. It is seen that the voltage gradient across the slice of GaAs is very high. Hence the electron velocity is also high, so that oscillations will occur at microwave frequencies. The Gunn effect is a bulk property of semiconductors and does not depend, as do other semiconductor effects, on either junction or contact properties. A threshold value of 3.3 kV/cm must be exceeded if oscillations are to take place. The frequency of the oscillations produced corresponded closely to the time that electrons would take to traverse such a slice of n−type material as a result of the voltage applied. This suggests that a bunch of electrons, here called a domain, is formed somehow, occurs once per cycle and arrives at the positive end of the slice to excite oscillations in the associated tuned circuit. When a voltage is applied, electrons flow as a current toward the positive end of the slice. The greater the potential across the slice, the higher the velocity with which the electrons move toward the positive end, and therefore the greater the current. Thus far, the device is behaving as a normal positive resistance. If the threshold value of 3.3 kV/cm is exceeded, so much energy is imparted to the electrons by the extremely high voltage gradient that instead of traveling faster and therefore constituting a larger current, their flow actually slows down. This is because such electrons have acquired enough energy to be transferred to the higher energy band, which is normally empty. This gives rise to the name transferred− electron effect, which is often given to this phenomenon. Electrons have thus been transferred from the conduction band to a higher energy band in which they are much less mobile, and thus the current has been reduced as a result of a voltage rise. If the applied voltage rises past the ”threshold negative−resistance value”, current falls, and thus the classical case of negative resistance is exhibited. HAM US/F-4 SaR Dec 2005 Domains Another phenomenon is the formation of domains. It is reasonable to expect that the density of the doping material is not completely uniform throughout our sample of gallium arsenide. Hence it is entirely possible that there will be a region, perhaps somewhere near the negative end, where the impurity concentration is less than average. In such an area there are fewer free electrons than in other areas, and therefore this region is less conductive than the others. As a result of this, there will be a greater than average potential across it. Thus, as the total applied voltage is increased, this region will be the first to have a voltage across it large enough to induce transfer of electrons to the higher energy band. In fact, such a region will have become a negative−resistance domain. A domain like this is obviously unstable. In fact, the whole domain moves across the slice toward the positive end with the same average velocity as the electrons before and after it, about 107 cm/s in practice. When it arrives at the positive end of the slice, a pulse is received by the associated tank circuit and shocks it into oscillations. It is actually this arrival of pulses at the anode, rather than the negative resistance proper, which is responsible for oscillations in Gunn diodes. When the domain arrives at the anode, there is once again sufficient potential to permit the formation of another domain somewhere near the cathode. It is seen that only one domain, or pulse, is formed per cycle of RF oscillations, and so energy is received by the tank circuit in correct phase to permit the oscillations to continue. Page 228 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 anode cathode For Training Purposes Only active zone Figure 113 HAM US/F-4 SaR Dec 2005 Radar Gunn Diode Page 229 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 IMPATT (impact avalanche and transit time) Diode A combination of delay involved in generating avalanche current multiplication, together with delay due to transit time through a drift space, provides the necessary 180 phase difference between applied voltage and the resulting current in an IMPATT diode. The cross section of the active region of this device is shown in Figure 114 . It is a diode, the junction being between the p+ and the n layers. An extremely high voltage gradient is applied to the IMPATT diode, of the order of 400 kV/cm, eventually resulting in a very high current. A normal diode would very quickly break down under these conditions, but the IMPATT diode is constructed so as to be able to withstand such conditions repeatedly. Such a high potential gradient, back−biasing the diode, causes a flow of minority carriers across the junction. If it is now assumed that oscillations exist, we may consider the effect of a positive swing of the RF voltage superimposed on top of the high dc voltage. Carrier velocity has now become so high that these carriers form additional carriers by knocking them out of the crystal structure, by so−called impact ionization. These additional carriers continue the process at the junction, and it now snowballs into an avalanche. If the original dc field was just at the threshold of allowing this situation to develop, this voltage will be exceeded during the whole of the positive RF cycle, and avalanche current multiplication will be taking place during this entire time. However, since it is a multiplication process, avalanche is not instantaneous. The process takes a time such that the current pulse maximum, at the junction, occurs at the instant when the RF voltage across the diode is zero and going negative. A 90 phase difference between voltage and current has thus been obtained. The current pulse in the IMPATT diode is situated at the junction. However, it does not stay there. Because of the reverse bias, the current pulse flows to the cathode, at a drift velocity dependent on the presence of the high dc field. The time taken by the pulse to reach the cathode depends on this velocity and of course on the thickness of the highly doped (n+) layer. The thickness of the drift region is cunningly selected so that the time taken for the current pulse to arrive at the cathode corresponds to a further 90 phase difference. Thus, as shown in the figure, when the current pulse actually arrives at the cathode terminal, the RF voltage there is at its negative peak. Accordingly, voltage and current in the IMPATT diode are 180 out of phase, and a dynamic RF negative resistance has been proved to exist. Such a negative resistance can be used in oscillators or amplifiers. Because of the short times involved, these can be microwave. The device thickness determines the transit time, to which the IMPATT diode is very sensitive. Accordingly, and unlike the Gunn diode, the IMPATT diode is essentially a narrowband device (especially when used in an amplifier). HAM US/F-4 SaR Dec 2005 Page 230 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 (a) Drift Region For Training Purposes Only (b) Harmonic of the current pulse at the cathode (a) applied DC plus RF voltage, (b) resulting current pulse and its drift across the diode Figure 114 HAM US/F-4 SaR Dec 2005 Radar IMPATT Diode Page 231 Part-66 Tunnel Diode The tunnel or Esaki diode is a thin-junction diode which, under low forward−bias conditions, exhibits negative resistance. This makes the tunnel diode, invented in the late 1950s, useful for oscillation or amplification. Because of the thin junction and short transit time, it is in use at microwave applications. The tunnel diode is a semiconductor p−n junction diode. It differs from the usual rectifier−type diodes in that the semiconductor materials are very heavily doped, perhaps as much as 1000 times more than in ordinary diodes. This heavy doping results in a junction which has a depletion layer that (with a typical thickness of 0.01 m) is so thin as to permit tunneling to occur. In addition, the thinness of the junction allows microwave operation of the diode because it considerably shortens the time taken by the carriers to cross the junction. A current−voltage characteristic for a typical germanium tunnel diode is shown in the figure. It is seen that at first forward current rises sharply as voltage is applied, where it would have risen slowly for an ordinary diode (whose characteristic is shown for comparison). Also, reverse current is much larger for comparable back bias than in other diodes, because of the thinness of the junction. The interesting portion of the characteristic begins at the point A on the curve ; this is the voltage peak. As the forward bias is increased past this point, the forward current drops and continues to drop until point B is reached; this is the valley voltage. At B the current starts to increase once again and does so very rapidly as bias is increased further. From this point the characteristic resembles that of an ordinary diode. The diode voltage−current characteristic illustrates two important properties of the tunnel diode. First it shows that the diode exhibits dynamic negative resistance between A and B and is therefore useful for oscillator and amplifier applications. Second since this negative resistance occurs when both the applied voltage and the resulting current are low, the tunnel diode is a relatively low−power device. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 232 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 A cathode B For Training Purposes Only anode characteristic of a tunnel diode Figure 115 HAM US/F-4 SaR Dec 2005 Radar Tunnel Diode Page 233 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 FREQUENCY DETERMINING ELEMENTS To build an oscillator, an amplifier is necessary to overcome losses in the circuitry, and an element that determines the frequency. The following elements may be used: S R-C-Wien bridges S L-C- resonant circuits S two wires parallel lines S micro strips on circuit boards S tank circuits S cavity resonators Only the last four are in use for microwave generation because it is not possible to produce separate inductors (L) or capacitors (C) for this frequency range. Two wires resonator behavior depends on its length related to the wavelength and the termination at the end of the lines. Three different situations are remarkable. The lines are 1. short circuit terminated 2. open 3. terminated with a resistor of the same value as the impedance of the line The last case simulates an infinite cable length and the principle is used for transmission lines to prevent reflections on the wire. Cases 1. and 2. can be applied to /4, /2 and other multiples of /4 lengths. A short circuit means always, that the current is a maximum and the voltage a minimum at this point. In the resonant case, it is easy to trace back on the line the behavior of voltage and current, to find out if the two wires are working similar as a serial- or parallel- resonant circuit. Example: A /4 line terminated with a short circuit works as a rejector circuit (parallel tuned), while if it is open at the end, it works as an acceptor circuit (serial). For /2 it is vice versa. HAM US/F-4 SaR Dec 2005 Page 234 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 For Training Purposes Only 4 8 4 & 4 2 8 2 How a two wires resonator with defined length reacts on different wavelengths. Figure 116 HAM US/F-4 SaR Dec 2005 Two Wires Resonator Page 235 Part-66 Tank Circuits and Cavities Rotating a two wires line, one will get a tank circuit. The advantage is that this will eliminate radiation losses. If also the middle conductor is removed, than a cvity resonator is achieved. higher capacitv For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Figure 117 HAM US/F-4 SaR Dec 2005 Tank Circuit Page 236 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 field distribution inductive coupling magnetic field lines resonance cavity electric field lines shorting piston capacitive coupling control voltage For Training Purposes Only resonance cavity varicap resonance cavity Figure 118 HAM US/F-4 SaR Dec 2005 Cavity Resonator Page 237 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 RADAR ANTENNAS There is a large number of antennas that are used for microwave applications, because most of the antennas can be used for this frequency range. One advantage of this frequency range is the small antenna dimension, if there have been no demands to the directivity. If high antenna gains and / or high antenna directivities are required, parabolic reflector - or flat plate - antennas are commonly used as airborne radar antennas. Two types of antennas are employed to obtain the required narrow beam, a directly fed parabolic reflector or a flat plate planar array. For a given diameter and wavelength, the flat plate has the higher gain/ least side lobe power. Since the flat plate is almost twice as efficient as the parabolic reflector it is used with a modern system. Parabolic Antenna The parabolic reflector works on a similar principle to a car headlamp reflector. Energy striking the reflector from a point source situated at the focus will produce a plane wave of uniform phase travelling in a direction parallel to the axis of the parabolic reflector. The feed in a weather radar parabolic antenna is usually a dipole. The Bendix parabolic antenna ANT-1T with a reflector diameter of 30’’ and for xband ( e.g. WXR at 9375 MHz ) has an antenna beamwidth of about 3 and is called a pencil beam. The antenna gain for such a pencil beam is 30 dB. The side lobes are 20 dB down ( main beam -20dB ). For ground mapping, there can be a spoiler grid integrated into the parabolic reflector dish. If this spoiler grid is activated, a cosec-diagram is developed. It is shaped to get ground echos. The activation of the spoiler grid is done by rotating the polarization from horizontal ( pencil beam ) to vertical ( cosec-diagram ). The rotation is done by applying a dc to a ferrite rotator within the circular waveguide feeder. If there is no MAP diagram, the antenna has to be tilted down to get ground returns. HAM US/F-4 SaR Dec 2005 Page 238 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Figure 119 HAM US/F-4 SaR Dec 2005 Parabolic Antenna Page 239 Part-66 Flat Plate Antenna The flat plate antenna or phased array antenna consists of strips of waveguide vertically mounted side by side with the road wall facing forward. Staggered off− centre vertical slots are cut in each waveguide so as to intercept the wall currents and hence radiate. Several wavelengths from the antenna surface, the energy from each of the slots will be summed in space, cancellation or reinforcement taking place depending on the relative phases. In this application the phase of the feed to each slot, and the spacing between slots, is arranged so as to give a resultant radiated pattern which is a narrow beam normal to the plane of the plate. The greater the number of slots the better the performance; since the spacing between the slots is depending on the wavelength we can only increase the number of slots by increasing the size of the flat plate. The Bendix flat plate antenna DAA-4A with an array diameter of 30’’ and for x-band has an antenna beamwidth of about 4. The antenna gain is 35 dB. The side lobes are 25 dB down ( main beam -25 dB ). For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 240 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 radiating slots RF power from transmitter For Training Purposes Only feeding wave guide wave guides with slots radiating slots Figure 120 HAM US/F-4 SaR Dec 2005 feeding wave guide Flat Plate Antenna Page 241 Part-66 Because of the limited size of the reflector or the flat plate, side lobes are generated leading eventually to nuisance targets. The flat plate has less side lobes than the parabola. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS HAM US/F-4 SaR Dec 2005 Page 242 Lufthansa Technical Training M13.4 NAVIGATION RADAR FUNDAMENTALS Part-66 ÅÅÅ ÅÅÅ ÅÅÅ ÅÅÅ ÅÅÅ ÅÅÅ nuisance cloud Side Lobes ÅÅÅ ÅÅÅ ÅÅÅ ÅÅÅ ÅÅÅ For Training Purposes Only cloud Figure 121 HAM US/F-4 SaR Dec 2005 Parabolic and Flat Plate Antenna Page 243 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Part-66 DME DME GENERAL The Distance Measuring Equipment (DME) is a Secondary Radar System, that measures the line−of−sight distance from the aircraft to a selected ground station. The DME ground stations very often are combined with VOR-, or ILS / MLS - stations. From the military TACAN- station the civil user can also use the DME part. Using a VOR/DME- or VORTAC- ground station, the PPOS can be calculated by the RHO / THETA navigation method. Using DME/ DME- or DME/ DME/ DME- ground stations, the PPOS can be calculated by the RHO / RHO or RHO / RHO/ RHO navigation method. The PPOS calculation is done by the FMS. The measured distance is sent for indication to the S navigation displays (NDs) and primary flight displays (PFDs) or/and S DDRMIs For Training Purposes Only The DME system operates in the L−band frequency range of about 1GHz.. Because of this frequency, the line of sight and therefor the maximum DME distance can be 320 nM. + Ǹ The DME interrogator is tuned to a VOR/ILS/MLS −DME paired ground station frequency S manually from a control panel or control display unit ( CDU ) or S automatically from a FMS. HAM US/F-4 SaR Dec 2005 Page 244 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Figure 122 HAM US/F-4 SaR Dec 2005 DME General Page 245 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION DME Part-66 DME PRINCIPLE The principle of DME navigation is based on transmission time measurements. The radiosignal needs 12.36 sec ( plus delay of the groundstation ), if the distance between DME interrogator and groundstation is one nautical mile. This time of 12.36 sec is called a Radar Mile. The selected DME groundstation DME or VOR / DME or VORTAC transmitts always a number of S 2.700 puls pairs / sec. These pp/s are all random pulses - SQUITTER -, if no aircraft with DME interrogator is in the vicinity of the groundstation. If an aircraft receives these squitter, the DME interrogator starts to be the activ part of the secondary radar system by interrogating the ground station. The ground station now reduces the number of squitter and transmits replysignals to the interrogation instead of squitter, so that the total number of pulses will be constant 2.700 ( squitter- and reply - puls pairs ). The ground station can handle approx. 100 aircraft. Pairs of interrogation pulses are sent at a random rate from an onboard interrogator to a DME ground station. After a defined delay (50 sec in X mode and 56 sec in Y mode) this ground station transmits the received signal by using another RF carrier.The difference between interrogation - and reply - frequency is always $ 63 MHz. The onboard interrogator receives this reply signal. The time between transmission and reception represents the distance or slant range between aircraft and ground station. The signals from the airborne equipment are paired pulses of a defined spacing ( 12sec or 36sec ), but the repetition rate varies from one transmission to the next. This principle - that is called JITTERING - allows the interrogator to find out its own puls pairs among many other pulses which can be replies for other aircraft or squitter. The ground station identifies itself by sending a morse signal at 1350 Hz. The interrogator receives from the ground station this Ident signal which is sent to the audio management unit for audio presentation in the flight deck and is decoded for Ident indication on the indicators. HAM US/F-4 SaR Dec 2005 Page 246 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Figure 123 HAM US/F-4 SaR Dec 2005 DME Principle Page 247 Lufthansa Technical Training M13.4 NAVIGATION DME Part-66 DME CHANNEL-FREQUENCY The DME frequencyrange is from 962 MHz up to 1213 MHz. The frequencies of the interrogation ( DL ) S vary from 1025 MHz to 1150 MHz, the frequencies of the reply ( UL ) S vary from 962 MHz to 1213 MHz. Transmit and receive frequency always differ by 63 MHz. Each transmit frequency has two paired receive frequencies one +63 MHz and one - 63MHz apart. The time gap between the two pulses is a function of the selected channel: S X-channel :12 s for interrogation( DL ) and reply ( UL ) S Y-channel : 36 s for interrogation ( DL ) 30 s for reply ( UL ) This defines X and Y channels. The channel spacing is 1 MHz. The total number of DME - Channels is 126 X- and 126 Y- Channels. The DME - frequencies or - channels are paired with VOR - or ILS / MLS - frequencies. FOR CHANNEL VOR - FREQUENCY 17 to 59 0.1 * ( CHANNEL NUMBER - 17 ) + 108 MHZ 70 to 126 0.1 * ( CHANNEL NUMBER - 70 ) + 112.3 MHZ DME - CHANNEL FOR VOR-FREQUENCY For Training Purposes Only 108 MHz to 112.2 MHz 112.3 MHz to 117.95 MHz HAM US/F-4 SaR 10 * ( VOR FREQUENCY - 108 ) + 17 10 * ( VOR FREQUENCY - 112.3 ) + 70 Dec 2005 Page 248 Lufthansa Technical Training M13.4 NAVIGATION DME Part-66 Channel Spacing is 1 MHz 36 s 30 s For Training Purposes Only 12 s X - Channels: 12 s 12 s pulse gap of interrogation pulse gap of reply Figure 124 HAM US/F-4 SaR Dec 2005 Y - Channels: 36 s 30 s pulse gap of interrogation pulse gap of reply DME Channel-Frequency Page 249 Part-66 Military Channels For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION DME Figure 125 HAM US/F-4 SaR Dec 2005 DME Channel-Frequency Chart # 1 Page 250 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION DME Figure 126 HAM US/F-4 SaR Dec 2005 DME Channel-Frequency Chart # 2 Page 251 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Part-66 DME SYSTEM ( Example A 320 ) Figure 127 Normally two independant DME systems are installed in the airplane. Each system consists of : S one DME interrogator S one DME antenna S one dual VOR/DME Radio Magnetic Indicator (VOR/DME RMI). For Training Purposes Only The following components can control the DME system: S for frequency/course selection S the Multipurpose Control and Display Unit 1(2) (MCDU), S the Radio Management Panel 1(2) (RMP) and S the Flight Management and Guidance Computer 1(2) (FMGC) S for audio controls S the CAPT (F/O) Audio Control Panel (ACP) and the Audio Management Unit (AMU) S for test causes S the MCDU and the Centralized Fault−Display Interface−Unit (CFDIU). The DME data are shown on : S the CAPT and F/O Primary Flight Displays (PFD) S the CAPT and F/O Navigation Displays (ND) S the VOR/DME RMI S the MCDU(s) (maintenance data). Frequency selection Controlling information ( for frequency/course selection ) is sent to the interrogators as an ARINC 429 control word from either the RMP or FMC( CDU ). This control word directs DME system operation by tuning the interrogator to a ground station and computing slant range distance information from that ground station. This distance is then provided to the units for display in the flight deck and to the FMC for flight navigation. HAM US/F-4 SaR Dec 2005 Suppression The DME / ATC and TCAS systems operate in the same L−band frequency range. To prevent interference between systems, only one system is permitted to transmit at any one time. A suppression circuit connects the DME, ATC and TCAS systems together and a mutual transmission suppression signal is provided from the on− line system to prevent the others from transmitting. Ident Each DME ground station periodically transmits an identification signal along with distance data.To identify the DME station, a 1350 Hz Morse code is transmitted. The higer audio frequency allows to distinguish the DME identifier from the collocated VOR identifier with 1020 Hz and the same three letter code. One of the identification signals of the foreground station is sent to the audio management unit for flight deck audio presentation. This system controls and directs the output to the headsets and/or the loudspeakers. The pilot can control the DME audio signals by pressing the VOR pushbutton switch on the ACP. In case of collocated ILS/ DME ground stations and when the ILS pushbutton switch is pressed on the EFIS control section of the FCU, the pilot can control the DME audio signal through the ILS pushbutton switch on the ACP. The identification signal is also decoded and then encoded by the interrogator into two ARINC 429, 32−bit data words. Each data word contains two characters of the ident display. The data words are sent to the DMCs for display on the NDs. The DME station ident will only be displayed when a VOR station ident is not available. The DME station ident can be identified, because it is displayed in smaller letters than the VOR ident display. Air/Ground The AIR/GROUND DISCRETE from the air/ground relay identify flight segement legs for the nonvolatile fault memory. It also inhibits the CMC from initiating a DME interrogator self−test when the airplane is in the air. Page 252 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Figure 127 HAM US/F-4 SaR Dec 2005 DME Syst. Schematic ( A 320 ) Page 253 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION DME Part-66 DME DATA DISPLAY S On the PFDs With ILS/DME collocated stations, the ILS/DME distance is shown in magenta in the L lower corner of the PFD (Item 1). These data come into view when you push the ILS pushbutton switch located on the EFIS control section of the Flight Control Unit (FCU). S On the NDs The VOR/DME distance is shown in green in the L lower corner of the ND for DME system 1 (Item 2), and in the R lower corner of the ND for DME system 2 (Item 3) when : − you set the mode selector switch on the EFIS control section of the FCU to ROSE (ILS, VOR, NAV) or ARC − you set the ADF/VOR/OFF switch to VOR. Each DME ground station periodically transmits an identification signal along with distance data. One of the identification signals of the foreground station is sent to the audio management unit for flight deck audio presentation. The identification signal is also decoded and then encoded by the interrogator into two ARINC 429, 32−bit data words. Each data word contains two characters of the ident display. The data words are sent for display on the NDs. The DME station ident will only be displayed when a VOR station ident is not available. The DME station ident can be identified because it is displayed in smaller letters than the VOR ident display.Should both VOR and DME idents be unavailable, the displays will indicate the current frequency (in MHz) the VOR is tuned to. In addition, when you push the VOR−D pushbutton switch on the EFIS control section of the FCU, this causes : Display of the VOR/DME and DME ground stations which are not already included in the flight plan, with the mode selector switch in ROSE NAV and ARC positions : S a circle for the DME- station S circle plus cross symbol for the VOR/DME- or VORTAC- station. HAM US/F-4 SaR Dec 2005 S On the VOR/DME RMI Two windows are available for indication of both distances from the DME 1 and DME 2 (Item 4) when the VOR/DME stations are collocated. When the DME or RMI monitoring circuits detect a fault, the corresponding display window (Item 5) is blanked. In case of No Computed Data (NCD) (out−of−range station) the windows show white horizontal dashed lines (Item 6). In addition, the LEDs on the face of the DME interrogator indicate the status of the DME system. Page 254 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Figure 128 HAM US/F-4 SaR Dec 2005 DME Data Display Page 255 Lufthansa Technical Training M13.4 NAVIGATION DME Part-66 DME MODES OF OPERATION DME system modes of operation are shown in the chart on Figure 129 . A description of the operation of each mode is given along with the flight deck display for that mode. The conditions for each mode are as follows: S STANDBY The STANDBY mode occurs at initial turn−on and is the fall−back mode when no ground station pulses are present. In this mode, the transmitter is off but the receiver counts the number of pulse pairs received from the ground station. S SEARCH When the DME is within range of the ground station, sufficient pulse pairs>650 will be counted by the receiver. This causes the DME to operate in the search mode. In the SEARCH mode, the interrogator transmits pulse pairs and looks for ground station replies. Reply pulses are stored in memory for later analysis. Upon reception of synchronous reply pulses, the DME converts the time to distance and switches to the track mode. S TRACK The TRACK mode occurs after lock−on in the previous search mode. Slant range distance is computed and provided to the Display Computer and FMC and is updated every 100 ms if there is no channel designated for display or 70 ms when designated by the FMC. The display is updated every 140 ms when the channel selected by the pilot. DME tolerance: S MEMORY The system operates in the MEMORY mode when station reply pulses are lost. The airplane range is then extrapolated from the most recent range data stored in memory. This occurs for approximately 10 seconds or, until sufficient signal strength is regained. S FAULT The DME enters the FAULT mode whenever it detects a fault. DME monitoring circuits continuously check the system and alert the flight crew if a fault has been detected by blanking DME displays. S SELF−TEST The SELF−TEST mode is initiated by pressing the test switch on the interrogator front panel. Indicators on the interrogator front panel show pass/fail and input status of the LRU. 0,1 nm up to a distance of 5 nm. For Training Purposes Only 0,2 nm between 5 nm and 200 nm. HAM US/F-4 SaR Dec 2005 Page 256 Lufthansa Technical Training M13.4 NAVIGATION DME Part-66 (RED) CONTROL INPUT FAIL (GREEN) PASS (RED) FAIL LR STATUS U MODE TEST STANDBY SEARCH DME INTERROGATOR TRACK SCAN Collins LRU STATUS For Training Purposes Only CONTROL FAIL DME−90 0 TEST Dec 2005 TRANSMISSION INHIBITED; RECEIVER (AND AUDIO) OPERATIVE >650 PULSE PAIRS PER SECOND RECEIVED TRANSMITTER ON AND OPERATING WITH HIGH PRF THREE−OR−MORE (OUT OF 5) CORRECT PULSE PAIRS RECEIVED INSEARCH MODE. MULTIPLE STATION TUNING FIVE STATIONS OR LESS SPECIFIED IN FOREGROUND LOOP UP TO 15 MAY BE SCANNED IN BACKGROUND LOOP MEMORY LOSS OF RETURN PULSES FOR 10 SEC COMPUTES RANGE (EXTRAPOLATED FROM PREVIOUS DISTANCE). LRU STATUS INDICATOR (RED/GREEN) FAULT CONTROL FAIL INDICATOR (RED) SELF− TEST INTERROGATOR MALFUNCTION (DURING CONTINUOUS MONITORINGOPERATION) PUSH AND HOLD THE TEST SWITCH, WAIT FOR THE INTERROGATOR TO CYCLE THROUGH THE SELF−TEST. INTERROGATOR FRONT PANEL LIGHTS WILL COME ON. RDMI DISPLAY 4 DASHES (NO COMPUTED DATA) 4 DASHES SLANT RANGE SELECTED STATION SLANT RANGE SLANT RANGE BLANK DISPLAY SEQUENCES THRU: 1. BLANK DISPLAY 2. DASHES 3. 0.0 NM SELF−TEST SWITCH Figure 129 HAM US/F-4 SaR DESCRIPTION OF OPERATION DME Modes Page 257 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Part-66 DME CONTROL The DME interrogator has two types of tuning control modes : S AUTOTUNE Mode Control of the DME interrogator is provided by either the left or right FMC. Based on information in the FMC data base or from information entered into the CDU, the DME interrogator gets a control word via an ARINC 429 data bus. This control word directs the DME system as required by the FMC. The NAV RADIO page on any CDU displays current DME system configuration tuning data and allows manual entries by the flight deck to change the existing DME tuning or control configuration. S MANUAL TUNE Mode The DME interrogator is controlled in the manual tune mode whenever the system does not have a valid input from the FMC. Control of the interrogator is then provided by the onside CDU or in Radio Nav Back-Up mode from a RMP. The left and right interrogators are therefore controlled exclusively by the left and right CDUs or RMPs respectively.The CDU or RMP can only tune the interrogator to only one station at a time. For Training Purposes Only DME SCANNING Modern DME Interrogator are called SCANNING - or AGILITY - DME. S DIRECTED SCAN In the DIRECTED SCAN mode, the DME has the capability of interrogating and providing distance information for one to five ground stations. − a)The first four directed channels (1, 2, 3, 4) are used for VOR paired DME. The fifth channel (5) is reserved for ILS paired DME ground station. Channels 3 and 4 are the primary, Channels 1 and 2 are the secondary VOR paired DME ground stations. The primary channels are predetermined by the FMC according to information in the flight deck data base. The secondary channels are used to obtain the best DME−DME pair for rho/rho navigation. The interrogator provides the FMC with frequency and binary (bnr) distance data at an update rate of 70 ms. − b)Only one of the five directed stations can be selected for display. When one directed channel is selected for display, the DME outputs a valid BCD distance word only for that channel. HAM US/F-4 SaR Dec 2005 S Three abnormal conditions exist, that will cause the DME not to operate in the one frequency DIRECTED SCAN mode. Under these conditions, the DME will automatically revert to FREE SCAN operation. The abnormal conditions are: S if the digital tuning signal falls below 5 words per second, S the tuning word sign/status matrix indicates an invalid condition, or S the word parity is incorrect. S FREE SCAN In FREE SCAN mode, the DME interrogator cycles through all 200 channels (excpt.the UHF paired DME channels ). The DME provides distance information for all stations that are within this range. The DME can independently scan all available channels which form a foreground loop and a background loop. a) Foreground loop The foreground loop consists of five stations, any or all of which can be designated by the FMC. If less than five stations are designated, the FMC fills these empty slot(s) with the closest DME paired station(s). The foreground loop stations are given priority service over the background loop stations. The display of the foreground stations are identical to that of the DIRECTED SCAN. The five foreground channels are restricted in use by the FMC as follows: - Channels 1 and 2 are reserved for the best DME−DME pair for radio updating. - Channels 3 and 4 are reserved for navaids manually (M) entered on by the CDU. These channels are also used in autoselected: route tuned stations (R), procedure tuned stations (P), or not M, R, or P autoselected stations (A), whenever manually selected stations have not be chosen. - Channel 5 is reserved exclusively for an ILS-MLS/DME paired station. The channel normally becomes active whenever the airplane is within 25 NM (direct distance) of the runway or within 10 NM of an approach leg. The FMC will autotune to a station for this channel only when the airplane is 40 NM or less from the runway. The FMC parks this channel whenever the above criteria is not met. b) Background loop The remaining 195 stations make up the background loop. The DME interrogator scans stations that respond with 450 squitter pp/s or more. These stations are then interrogated, distance information is calculated and information is sent to the FMC. Distance information is updated every 2 seconds. Page 258 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Part-66 TUNING MODE: A = AUTOTUNE M = MANUAL TUNING R = ROUTE TUNING P = PROCEDURE TUNING LEFT VOR/DME STATION Radio Management Panel RMP NAV V O R C R S MLS/DME OR ILS/DME STATION V O R A SEA 1 84 172 L −−− P AR K P R E S E L E C T 1 0 9 . 6 0 / 2 3 7 For Training Purposes Only C R S A D F R − − − − − − 1304. 5 B F O I L S − ML S < 1 1 0 . 7 0 / 1 2 8 ° R E L NM1 1 7 . 9 0 R A D I A L −−− A D F R ADI O L 116. 80 RIGHT VOR/DME STATION PRESELECTED STATION −−−−−− PRESELECT PROMPT NAV RADIO CDU Page Figure 130 HAM US/F-4 SaR Dec 2005 DME Control Page 259 Part-66 DME ANTENNA The L band antenna is a short stub, all aluminum, blade−type antenna which operates in the 960 MHz frequency band. It is vertically polarized, has an impedance of 50 Ohm and a VSWR of 1.42 : 1. The antenna is grounded on the RF center conductor for use in self−test circuits. NOTE:The bonding contact resistance between the antenna (3) and the fuselage structure must be less than 5 milliohms. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME HAM US/F-4 SaR Dec 2005 Page 260 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION DME Figure 131 HAM US/F-4 SaR Dec 2005 DME Antenna Page 261 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Part-66 LRRA LRRA GENERAL The Low Range Radio Altimeter ( LRRA or RA ) is a Primary Radar System, that measures the Height between the extended MLG and the Ground - Height Above Ground Level ( AGL ) -. The operational „altitude“ is - 20 ft to + 2.500 ft. One of the main characteristics of the system is that it locks onto the leading edge of the reflected wave. This permits to measure the distance between the aircraft and the nearest obstacle. The radio altimeter can therefore operate over non−flat ground surface. The system operates between 4.2 GHz and 4.4 GHz ( C- Band ). LRRA Principles: Three different methods of measurement are in use: For Training Purposes Only S Pulse modulated RA S FM-CW RA with constant Modulation Frequency S FM-CW RA with constant Difference Frequency Definitions: S Altitude is the barometric altitude of the aircraft above MSL. QNH ( actual pressure) is selected in the baro set window of the altimeter. Indication on the altimeter. S Elevation is the vertical distance between MSL and fixed points on SFC (surface), e.g tower, airfield etc. S Flight Level is the barometric altitude of the aircraft above standard pressure level of 1013.25 hPa (Value set in the baro window of the altimeter). S Height is the vertical distance between Ground and aircraft. Indication is at the Radio (height) Altimeter RA. HAM US/F-4 SaR Dec 2005 Page 262 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Figure 132 HAM US/F-4 SaR Dec 2005 RF Frequency : 4300 MHz (4.3 GHz) Radiated Power : 5 Watt peak (pulse) or 100 mWatt (FM-CW) Operational Height : - 20 ft to 2.500 ft above Ground Accuracy : 3% or 3 ft LRRA Gweneral Page 263 Part-66 LRRA GENERAL DESCRIPTION The radio altimeter system has normally two receiver/transmitters (R/Ts). Each R/T has a transmit and a receive antenna. The radio altimeter R/T generates a RF signal that is sent to the ground and reflected back to the airplane. The time that it takes for the signal to travel from the transmit circuit of the R/T to the receive circuit of the R/T is changed into absolute altitude. The number one system altitude is shown on the captain display and the number two system shown on the first officer display. The altitude data and signal validity is sent as an analog signal or on two ARINC 429 data buses. The ARINC 429 data buses send data to these components: S display electronic units (DEU) or display management computer (DMC) S Flight control computers (FCC) S Ground proximity warning computer (EGPWC) S Flight Warning Computer (FWC) S Traffic alert and collision avoidance system (TCAS) computer S Weather Radar (WXR) For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA HAM US/F-4 SaR Dec 2005 Page 264 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Part-66 TRANSMIT ANTENNA RECEIVE ANTENNA FCC FWC COMPUTER AUTO CALLOUTS EGPWC TCAS COMPUTER For Training Purposes Only RA RECEIVER/ TRANSMITTER (2) CFDIU DEU/DMC WEATHER RADAR Figure 133 HAM US/F-4 SaR Dec 2005 LRRA General System Schematic Page 265 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 LRRA PRINCIPLE The principle is based upon the measurement of the time , which a radiosignal needs to travel from the aircraft down to the ground and back. The velocity v of radiowaves in free space is: S v = C = 3 x 10 8 m/s = 300.000 km/s S v = C = 9,836 x 10 8 ft/s The time is called and is a function of the height: t+ ( e.g. Honeywell RT-300 ) A 20 kHz modulator generates a 100− nanosecond modulation pulse every 50 s. A jitter generator modulates the modulator pulse repetition frequency (PRF), to prevent display of second time around lock- on at high altitudes. A master- oscillator is generating a 4.3 GHz RF- signal. The 100 nsec pulse modulates the 4.3 GHz. After amplification this pulsed RF signal of nominal 5 watts peak power is send via the transmit antenna to the ground. A sample of the transmitted signal is detected and applied to a processor as a Time Zero T0 reference. The RF- ground return signal from the receiver antenna is amplified and applied to the processor. Within the processor the received signal is compared to the sync. pulse T0. A timedifference of 2 nanoseconds represents one foot Radio Height. The processed time is converted to a dc analog voltage or an ARINC data bus. For Training Purposes Only The Certification for LRRAs are determined in S TSO C - 87 and S RTCA DO - 160A Pulse Transmission Principle HAM US/F-4 SaR Dec 2005 Page 266 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 HEIGHT OUTPUT CONVERTER T= 2 nsec / ft T0 Pulse HEIGHT H PULSED TRANSMITTER PROCESSOR Video RECEIVER For Training Purposes Only 4.3 GHz Figure 134 HAM US/F-4 SaR Dec 2005 LRRA Pulse principle Page 267 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 FM/CW Const. Modulation Freq. Principle ( e.g. Collins LRA-700 ) The transmitter sends a frequency modulated ( FM ) 4.3 GHz CW signal to the ground and the receiver receives the reflected signal. The modulation is done by a constant 100 Hz triangle signal. The periodtime T is a constant of 10 msec. Due to the FM, the transmit frequency changes between 4,25 GHz and 4,35 GHz. The frequency deviation F is a constant of 100 MHz. During travel time of the electromagnetic wave, after leaving the transmitter, it does not change its frequency anymore. In the mean time inside the transmitter, the frequency has changed due to the FM. The receiver mixes the received and the actual frequency and the result is a frequency difference f. Because T, F and c are constant values, this frequency difference f - called beat frequency - is proportional for the travel time and therefore also for the height above ground. At a height of 2.500 ft, the frequency difference f will be about 100 kHz. t+ D D + t ń D + D ń +D D For Training Purposes Only XD HAM US/F-4 SaR Dec 2005 Page 268 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 ËËË ËËË ËËË ËËË ËËË ËËË F ( GHz ) LOW Height 4,35 DF F ( GHz ) 4,3 4,35 4,25 T DF t 0 4,3 5 Df F ( GHz ) For Training Purposes Only DF 4,3 T X X T t ( msec) 0 5 Dec 2005 t 0 5 10 ( msec) 10 Figure 135 HAM US/F-4 SaR 4,25 ËË ËË ËË ËË ËË ËË ËË ËË HIGH Height 4,35 4,25 10 ( msec) LRRA with Constant Modulation Period Page 269 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION LRRA Part-66 FM/CW Const. Difference Freq. Principle ( e.g. Thomson-CSF ERT-530 ) The transmitter sends a frequency modulated ( FM ) 4.3 GHz CW signal to the ground and the receiver receives the reflected signal. The modulation is done by a variable sawtooth signal. Due to the FM, the transmit frequency changes between 4,24 GHz and 4,36 GHz.The frequency deviation F is a constant of 123 MHz. The operating mode of the RA transceiver is based on the leading−edge tracking FM/CW (Frequency Modulation/Continuous Wave) principle: S the FM/CW signal is transmitted towards the ground, reflected, and received after a delay ( ) depending upon the aircraft height above the ground ( and the AID ). S Because the transmitted frequency is continuously changing, at any instant a difference will exist between the transmitted frequency and the received frequency. S In the RA, the sweep period T of the sawtooth is varied, and the difference or beat frequency fb is held constant at 25 kHz. This is done, by driving the VCO from the sawtooth generator with a constant amplitude but variable frequency. S A monitor discriminator checks, that the beat frequency spectrum is centered on fbo ( fbo is the desired value of 25 kHz ). S In the absence of this condition, a SEARCH cycle is triggered. S During SEARCH , the sawtooth period is varied from lowest to highest limit. One SEARCH cycle takes about 0.3 sec. S If the beat frequency fb is 25 kHz, the system switches to TRACK, and the height value is transmitted to the user. S The SEARCH / TRACK mode is done three times a second. S In TRACK, the Height information is derived from the sawtooth period time T. This is the only variable parameter proportional to height and is converted to S digital ARINC 429 and nonlinear dc outputs. HAM US/F-4 SaR Dec 2005 %" D ' %'"!& D + t ! t+ $('"! "% ' ' & )! , + D D Page 270 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 F ( GHz ) F ( GHz ) 4,36 4,36 4,3 LOW Height DF 25 kHz Df = const 25 kHz 4,24 4,3 X X T ' F ( GHz ) 4,36 4,24 For Training Purposes Only X X 25 kHz T HIGH Height 4,3 t 4,24 X X Figure 136 HAM US/F-4 SaR Dec 2005 T ' LRRA with Constant Difference Frequency Page 271 Part-66 LRRA Search / Track ( LEERER MERKER) S Calibration and Search Graph A is the CLOCK generator output and is a series of 50msec pulses, 300msec apart. Calibration and search occurs before the system locks on and displays correct radio altitude. Internal measurements occur in the first 50msec. The circulators connect the transmitter and receiver to the internal delay line, loop errors are detected and compensated, and the altitude scan limit is set. During the following 300msec, two complete cycles of 0−5000 ft. scans occur; during this time the frequency of the sawtooth modulating signal to the transmitter changes and the circulators connect the transmitter and receiver by the antennas. Initial warmup takes 10 seconds, so the system does not transfer to TRACK operation at the first occasion that the height loop could lock on. Graph B shows the sawtooth waveform variation and Graph C shows the Self−Calibration ENABLE conditions. Note that self calibration is inhibited for the first 20msec of the internal checks, this is because the receiver is not tuned to FT at this time; but it is enabled for the last 30msec of internal checks. Graph D shows TRACK mode is enabled during internal checks, this permits the integrator to correct loop gains, using the internal reference delay. S Tracking Height less than 5000 ft Graph B sawtooth is the same as in SEARCH for the first 50 msec. During internal checks. During Self−Calibration (Graph C), the loop correction keeps the slope (modulation period) constant and it corresponds to 0 ft. Self−Calibration is inhibited for 20 msec, after each ENABLE period, during this time the sawtooth generator slope progressively changes in SEARCH mode. (Graph D). Graph E shows when the periodmeter stores and calculates a DC voltage related to height, this voltage varies according to a semilogarithmic law (linear below 480 ft). Height evaluation is achieved by measuring the modulation sawtooth duration, and is recalculated three times per second. Periodmeter storage is disabled at the end of the first sawtooth having the correct slope. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION LRRA HAM US/F-4 SaR Dec 2005 Page 272 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION LRRA Figure 137 HAM US/F-4 SaR Dec 2005 Tracking height less than 5000 ft Page 273 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 LRRA SYSTEM FAILURES Performance Certification The LRRA meets the requirements of the TSO C-87 and RTCA DO-160. According to the TSO, the following tolerances are valid for pitch up to 15 and bank up to 20: Height ( ft ) Vertical Speed ( ft/sec ) Accuracy 3 - 100 0 - 15 3 ft 100 - 500 0 - 20 3% 500 to max. 0 - 20 5% Turn- Around- Time This failure happens only at the constant modulation principle. At each period there are two times ( Turn around Times TT ), in which the relation between Height and difference frequency is not proportional. For a height of 1.000 ft the TT is about 2 sec. Within one period there are two TT, so total TT is about 4 sec per period. The time for one period is 0.01 sec or 10.000 sec. So the margin of error is 0.04% and will decrease at lower heights. Doppler effect The equation D +t does not take into account the doppler effect. In fact when the vertical speed of the aircraft is not zero, the equation is: ǒ Dń Ǔ ) ǒńlǓ where l is the doppler frequency, is the vertical speed and l is the signal D +t wavelength. For a sinkrate of e.g. = 20 ft/sec and an average frequency of 4.300 MHz, the dopplershift is about 175 Hz. FM/CW Const. Modulation Freq. Principle The doppler effect increases with lower heights, because the difference frequency is low at low heights ( e.g. circa 4 kHz at 100 ft and circa 100 kHz at 2.500 ft ) and the amount of dopplershift does not change with height, if the vertical speed is constant. At 2.500 ft the error is about 0,17% and at 100 ft about 4.37% for a sinkrate of 20 ft/sec. For this principle, the doppler effect can be reduced by averaging the difference frequencies f1 and f2 . D + For Training Purposes Only ǒ Dń Ǔ D ) D FM/CW Const. Difference Freq. Principle Because the difference frequency is constant at 25 kHz and the amount of dopplershift does not change with height, if the vertical speed is constant, the doppler effect is independent from height. For all heights, the error is about 0.7 % at a sinkrate of 20 ft/sec. At climb, the measured T is too high and at descent it is too low. For this principle, the doppler effect can be reduced by an appropriate delay in the measuring circuit. HAM US/F-4 SaR Dec 2005 Page 274 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 F ( GHz ) F ( GHz ) 2fd Df2 4,35 4,35 4,3 Df1 4,25 4,25 TT X X 0 OT TT 5 OT TT 10 t ( msec) Descent 2fd 4,36 4,3 OT : Operational Time t F ( GHz ) TT : Turn around Time For Training Purposes Only T X X Df = const 25 kHz 4,24 Figure 138 HAM US/F-4 SaR Dec 2005 X X T t LRRA System Failures Page 275 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION LRRA Part-66 LRRA SYSTEM ( Example A 320 ) Figure 139 Normally the radio altimeter comprises two independent systems. Each system consists of : S one transceiver S one transmission antenna S one reception antenna S one fan. In addition, the Centralized Fault−Display Interface−Unit (CFDIU) controls the system through the Multipurpose Control and Display Unit 1 (2) (MCDU) for test causes. Aircraft Installation Delay ( AID ) The (AID) program pin is grounded to the 57 feet selection. This calibrates the system, so that the radio altitude is 0 feet at touchdown at a normal landing configuration. It makes an allowance for these conditions: S Length of the antenna cables S Fuselage to ground distance S Flare angle. System Select Program Pins The system select program pin input sets: S the system identification and S synchronization between system # 1/2/3. If more than one RA System is installed, a synchronization is necessary to prevent a mutual influence. The synchronization can be done by : S different fixed modulation rates for the systems (e.g. RA # 1:145Hz, RA # 2: 155Hz). S different variable modulation rates for the systems, by using jitter generators. S phasing cable between the systems, so that the modulation frequencies are synchronized to fixed phases ( 0, 120, 240 ). HAM US/F-4 SaR Dec 2005 Page 276 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION LRRA Figure 139 HAM US/F-4 SaR Dec 2005 LRRA Syst. Schematic ( A 320 ) Page 277 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Part-66 LRRA DATA DISPLAY The radio height data is shown on the Primary Flight Display (PFD). In normal operation, system 1 provides information to the CAPT PFD and system 2 to the F/O PFD. In non EFIS aircraft, the indications can be performed on separate LRRA indicators. S On the PFDs 1. Digital Height data display (Item 3) The aircraft height data with respect to the ground is shown in the sector 2 of the PFD. This indication is at the bottom of the attitude sphere for height less than or equal to 2500 ft. The dimension and color of the digits change in relation to the height (H) and decision height (DH) as follows: H > or = to 400 ft.................. 3 mm green digits 400 ft > H > DH + 100 ft ..... 4 mm green digits H < DH + 100 ft .................... 4 mm amber digits The resolution of the display is also a function of the height: H > 50 ft ................................ 50 ft > H > 5 ft .................... H < 5 ft .................................. 10 ft increments 5 ft increments 1 ft increments For Training Purposes Only NOTE: If no DH has been set by the pilot: H < 400 ft: ............................ 4 mm amber digits. 2. Analog Height data display (Item 2) − When the aircraft is below 570 ft height above the terrain: 3. Decision height display (DH) (Item 1) The DH data are shown on the right top corner of the PFD as soon as the radio altimeter operates. When the height is lower than the DH, a DH amber warning message comes into view at the bottom of the attitude sphere (Item 5). 4. Failure indication With failure of both radio altimeters,: S a red RA warning message appears in place of the digits, if in slats extended configuration S the ribbon goes out of view and S the limit of sector 2 remains at its lower position. S On LRRA Indicators Indications can be performed on separate low range radio altimeter indicators and/or in the attitude director indicator. The LRRA indicator displays radio altitude from a range of minus 10 feet to plus 2500 feet by means of the altimeter pointer. The DH selector knob rotates the DH cursor around the periphery of the indicator. When the actual height is equal to or less than the decision height a signal is provided to illuminate a Decision Height Light. Above 2500 feet the altimeter pointer is biased behind the mask. The TEST switch initiates a LRRA system test, provided conditions are suitable. a red ribbon comes into view on the bottom and at the right of the altitude scale and moves up as the aircraft is in the descent phase (Item 2). When the aircraft has touched down the ground: the top of this ribbon is at the middle of the altitude window. − When the aircraft is below 150 ft, the height is shown by the distance between the horizon line and the limit of the sector 2. The limit of the sector 2 moves up as the aircraft is in the descent phase. The distance between these two lines is proportional to the ground height. As it moves up, the limit line erases the graduations on the pitch scale. HAM US/F-4 SaR Dec 2005 Page 278 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Part-66 POINTER MASK POINTER WARNING FLAG DH CURSOR SELF TEST SWITCH DH CURSOR CONTROL For Training Purposes Only Moving Height Tape: 0 - 2.500 ft Figure 140 HAM US/F-4 SaR Dec 2005 LRRA Indicators Page 279 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 LRRA OUTPUTS The Radio Height Datas are transmitted in compliance with the S low speed ARINC 429 standard or S ARINC 552 as an analog DC voltage. Additional to the height proportional outputs, there can be discrete outputs at specific heights ( Altitude Trips ). Digital Output ARINC 429 ARINC 429 transmits RA Height with S label 164 as BNR Height S label 165 as BCD Height S label 356 as Height Maintenance Status For Training Purposes Only Analog Output ARINC 552 Elder RAs transmit the height as DC voltage. Up to a of+480 ft , the relation between height and outputvoltage is linear, and for more than + 480 ft the relation is logarithmic. HAM US/F-4 SaR Dec 2005 Page 280 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION LRRA Part-66 30 H(ft) −20 0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400 2500 VOLT 0.000 0.400 2.400 4.400 6.400 8.400 10.392 12.151 13.646 14.947 16.098 17.130 18.065 18.920 19.708 20.438 21.119 21.756 22.355 22.920 23.455 23.962 24.446 24.907 25.347 25.769 26.174 BELOW 480 ft: 25 U = 0.02H + 0.4 20 ABOVE 480 ft: e x ( H + 20 ) U = 10 ln 500 15 10 5 0 BELOW 500 ABOVE 1000 1500 2000 2500 HEIGHT (FEET) GROUND LEVEL Figure 141 HAM US/F-4 SaR Dec 2005 LRRA Analog Output Standard Page 281 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Part-66 LRRA ANTENNAS The transmission- and reception- antennas are identical. Two types of antennas are in use: S horn radiator type and S microstrip type. For Training Purposes Only Horn radiator type Each antenna is in the form of a truncated pyramid. The antenna is attached to a flat circular base which includes a locating pin and is installed flush with the aircraft structure. The antenna is supplied through a coaxial connector linked to the transceiver. HAM US/F-4 SaR Dec 2005 Microstrip type The transmission and reception antennas are identical and of the type ”microstrip”. The small thickness microstrip antenna is installed on the skin of the aircraft and fixed by four screws, sealing is by an O−ring. Each antenna is supplied through a coaxial connector linked to the transceiver. The operating range of the antenna according to the attitude of the aircraft is limited to + or − 30 for roll and pitch angles. Installation of the RA-Antennas For an accurate function of the system, the correct installation of the antennas is very important. The distance between Xmit- and Rcv- Antenna must be large enough to get a sufficient decoupling ( > 40 dB ). But the distance must be small enough, to get the necessary AntennadiagramOverlapping at low heights ( typical distances are 0.5 to 1.0 m ). For installation, put the antenna (4) into the correct position and install the screws (5) to the fuselage structure. Make sure that the FWD sign or the datum mark on the antenna points in the direction of flight. The antenna mounting screws (5) give the electrical bonding between the antenna (4) and the fuselage structure (1). So make sure that the antenna−attachment screw−head recesses have no signs of paint, corrosion or sealant. Make sure that the bonding contact resistance between the antenna (4) and the fuselage structure (1) is not higher than 5 milliohms. Make the contour of the sealant on the screws (5) and around the antenna (4) smooth. Page 282 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION LRRA Part-66 XMIT#2 RCV#2 RCV#1 For Training Purposes Only XMIT#1 Horn Type Microstrip Type Figure 142 HAM US/F-4 SaR Dec 2005 LRRA Antennas Page 283 Lufthansa Technical Training M13.4 NAVIGATION LRRA Part-66 LRRA TESTS For trouble shooting, there are three tests that can be performed: S OPERATIONAL Test S BITE or ARINC Test S RAMP Test OPERATIONAL Test This test is done by activating all RA Systems ( PWR ON ) and to check the Height indications on the indicators. Check the transfer- and warning- functions, if one system is switched off. BITE or ARINC Test ( see Figure 143 ) This test switches off the antennas and activates a test delay-line within the LRRA. The test indication depends on the test delay-line and is normally : 40 ft. NOTE: During the test of syst. #1, you can hear the Automatic Call Out announcements. You must open the RA 1 circuit breaker to hear the Automatic Call Out announcements during the RA 2 test. For Training Purposes Only RAMP Test (see Figure 144 ) Reason for the Job is, to do the automatic simulation of a change in altitude from 500 ft to 0 ft. For this example ( A 340 ) two successive ramps are available: S RAMP1: 500−52 FT (10 FT/S) S RAMP2: 52− 0 FT ( 3 FT/S) Trouble shooting ( see Figure 143 and Figure 145 ) For additional TS, REPORT - and TROUBLE SHOOT- Pages can be used. HAM US/F-4 SaR Dec 2005 Page 284 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION LRRA Figure 143 HAM US/F-4 SaR Dec 2005 LRRA Testing Page 285 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION LRRA Figure 144 HAM US/F-4 SaR Dec 2005 LRRA Testing Page 286 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION LRRA Figure 145 HAM US/F-4 SaR Dec 2005 LRRA Testing Page 287 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Part-66 ATC TRANSPONDER GENERAL The Air Traffic Control (ATC) system ( SSR-Secondary Surveillance Radar System ) is based on the replies provided by the airborne transponders in response to interrogations from the ATC secondary radar ground station. The ground ATC secondary radar uses technics which provide the air traffic control with information that cannot be acquired by the primary radar. This system enables to distinguish between aircraft and to maintain effective ground surveillance of the air traffic. The system provides the air traffic controllers with : −Mode A : transmission of aircraft identification or, −Mode C : transmission of aircraft barometric altitude or, −Mode S : aircraft selection and transmission of flight data for the ground surveillance. The mode S is fully compatible with the other modes, A and C. The mode S bas been designed as an evolutionary addition to the ATC system to provide the enhanced surveillance and communication capability required for air traffic control automation. For Training Purposes Only NOTE:The ATC / Mode S will be able to provide the Traffic Collision Avoidance System (TCAS) with the aircraft address. HAM US/F-4 SaR Dec 2005 Page 288 Part-66 THIS PAGE INTENTIONALLY LEFT BLANK For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 289 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Part-66 PRINCIPLE OF OPERATION The air traffic control (ATC) system contains airborne components required to enable ground facilities to track airplane movement through ground facility sectors. The ground facilities monitor the airplane’s location, direction of travel, identification and altitude above 1013.25 hPa. The airplane transponder responds, with a reply, to an interrogation from the ground station. The reply signal contains the coded information required by the ground facilities. The ATC system receives interrogation signals from the ground station at a frequency of 1030 Mc via the antenna and routes the signal to the transponder. The transponder processes the signal and returns a coded reply signal at a frequency of 1090 Mc back to the ground station. The reply signal is coded by signals, depending on the mode of operation. The ATC and DME systems are both in the same frequency range. To prevent interference between receiver transmitter (RT) only one RT can transmit at any instant. Suppression signals prevent simultaneous transmission. Secondary Surveillance Radar System ( SSR ) To overcome these weaknesses, the ATC system also uses a secondary surveillance radar system. This system is the basis of the ATCRBS. The secondary surveillance radar, whose antenna also rotates (and is usually co−located with the primary radar antenna), transmits an interrogation signal to transponder−equipped aircraft within range. This triggers a transponder reply signal which is received by the secondary radar. Since the reply signal is not an echo of the secondary radar pulse, it is not subject to the same weakening effects as that of the primary radar echo signals. Additionally, the reply signal can be encoded with pulse combinations. This allows each transponder−equipped aircraft to be assigned a unique reply signal and also permits encoded altitude information (Mode C operation) to be transmitted in the reply information. Thus, the ATC controller can know the range, azimuth, altitude, and unique identification of each transponder−equipped aircraft within his radar range. For Training Purposes Only Ground stations The Air Traffic Control system uses two types of radar: S primary surveillance and S secondary surveillance. Primary Surveillance Radar System ( PSR ) The primary surveillance radar is a conventional radar system that transmits pulses of rf energy and listens between transmissions for an echo of the energy from targets within range. The echo signals are processed and displayed on a radarscope. The distance to the target is determined by timing the period between the transmision of the pulse and its received echo. The direction of the target’s determined by knowing the direction in which the antenna is pointing when the echo is received. By synchronizing the scan of the radarscope display with the time of the transmitted pulses and the direction of radar antenna orientation, target range and azimuth are determined. The primary surveillance radar has several weaknesses. The signal, normally line−of−sight, can be distorted by certain weather conditions such as temperature inversions, and the echo returned by small targets may not be of sufficient strength to give reliable signals on the radarscope display. HAM US/F-4 SaR Dec 2005 Page 290 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 146 HAM US/F-4 SaR Dec 2005 Air Traffic Control Page 291 Part-66 Ground Radar Display The secondary surveillance radar received signals are processed similar to those of the primary radar received signals and displayed on a radarscope. At ATC facilities having both types of radar systems, the two signals are superimposed display on the same radarscope. This allows the ATC controller to observe all aircraft (within range), whether or not they are transponder equipped. At some locations within the ATC system, only secondary surveillance radar systems are available. These fill the gaps between primary surveillance radar sites, but limit the information displayed to the controller to just that received from aircraft transponders. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 292 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Part-66 LN 123 140 46 For Training Purposes Only X X X X Figure 147 HAM US/F-4 SaR Dec 2005 F H I,J L Weather Radar Lines ATC Emergency Signal Direction Finder Lines SSR Target- Tracked, Data Tag norm. Pos. M O P Q SSR SSR SSR SSR R PSR (Primary Surveillance Radar) Target Target- Tracked, Data Tag switched Pos. Target- not tracked, Data Tag norm. Target- non tracked, no Data Tag. Target- not identified ATC Ground Radar Scope Page 293 Part-66 Principle of interrogation An airborne transponder provides coded reply signals in response to interrogation signals from the ground secondary radar and from aircraft which will be eventually equipped with the TCAS. This ground interrogation is transmitted in the form of: S pulses P1 and P3 for the mode A or C ( ATCRBS ) or S pulses P1, P3 and P4 for the mode S all call or S pulses P1, P2, P5 and P6 for the mode S selectiv call. Each ground interrogator transmits its interrogations at the frequency of 1030 MHz in the form of a series of pulses. Depending on the pulse intervals and numbers, they define four different interrogation modes: −Mode A interrogations (identification) −Mode C interrogations (altitude range) −Mode S all call interrogations −Mode S interrogations. After receiving these pulses, the transponder identifies and decodes the interrogations. Depending on the detected interrogation mode, the transponder transmits the identification of the aircraft or its barometric altitude. The transponder is provided with the Mode S Data Link capability. The transmission of the replies to the ground takes place on a carrier frequency of 1090 MHz. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 294 Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Part-66 For Training Purposes Only or 21 s Figure 148 HAM US/F-4 SaR Dec 2005 ATC Interrogation Principle Page 295 Part-66 ATCRBS INTERROGATIONS ATCRBS interrogations consists of two pulses labelled P1 and P3. These pulses are transmitted by the directional antenna of the Secondary Surveillance Radar (SSR) and spaced according to the SSR mode of operation. S Modes 1 and 2 are for military use only. S The military mode 3 and the civil modes A and B are interrogations, to get the selected code from the ATC control Panel ( alpha squawk ). S The mode C interrogation is asking for the coded barometric altitude . Coded altitude is taken out of pilots altimeter or airdata computer in increments of 100 ft based on 1013.2 mb. S The mode D is not jet defined. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 296 Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Part-66 MODE 1 2 For Training Purposes Only 3/A UTILIZATION PULSE INTERVAL in sec s military P1 s military P1 military and civil ATC P1 P3 s civil ATC C civil ( altitude coded ) P1 P3 s P1 P3 s civil ( not jet defined ) Figure 149 HAM US/F-4 SaR P3 s B D P3 Dec 2005 P1 P3 ATCRBS Interrogation Codes Page 297 Part-66 ATCRBS Side Lobe Suppression SLS A control pulse P2 is radiated by the omnidirectional antenna of the SSR two microseconds after the pulse P1 is transmitted from the directional antenna of the SSR. Amplitude of pulses P1 and P3 varies according to the position of the directional antenna. P1 and P3 are at higher amplitude than P2 when the aircraft flies in the antenna main beam. Amplitude of the pulse P2, also called side lobe suppression pulse SLS, is constant whichever the position of the omnidirectional antenna. The aircraft receives the P1 pulse at a higher amplitude than P2. The aircraft transponder detects this amplitude difference and determines the interrogation to be a valid interrogation. When the transponder detects that P2 is within 9 dB of P1, in amplitude, it determines that the interrogation is not valid ( Side lobe interroation ). For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 298 Part-66 ATCRBS TRANSPONDER REPLIES For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER ATCRBS TRANSPONDER DOES NOT REPLY Figure 150 HAM US/F-4 SaR Dec 2005 ATCRBS Side Lobe Suppression Page 299 Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Part-66 MODE S INTERROGATIONS The ATC/Mode S transponder equipped aircraft and ground station enhance the operation of the ATCRBS system by adding: S a data link capability, S discrete interrogation capability, and S performance improvements. In addition, the ATC/Mode S transponder is still capable of operating with an ATCRBS only ground station. The airborne equipment remains the same as the airborne ATCRBS equipment except the ATC transponder is replaced with an ATC/Mode S transponder. S ATCRBS / Mode S all call interrogation If the received interrogation contains a S 1.6−microsecond pulse in the P4 position and the amplitude of S P4 is above the amplitude of P3 less 1 dB, a mode S reply is generated 128 microseconds after the leading edge of the pulse P4. ATCRBS / Mode S all call interrogation A mode ATCRBS / Mode S all call interrogation consists of three pulses labelled: P1, P3 and P4. They are transmitted by a directional antenna. A control pulse P2 (Side Lobe Suppression Pulse) is transmitted following P1 by an omnidirectional antenna. The detection of pulse P1, P2 and P3 is the same as for mode A or C interrogation. For Training Purposes Only The Mode S Transponder will detect the interrogation as an: S ATCRBS only interrogation If the width of the pulse S P4 is 0.8−microseconds ( mode A/C only all−call ), and the amplitude of S P4 is above P3 minus 1 dB, no reply from a mode S transponder is transmitted. S ATCRBS interrogation If the the amplitude of S P4 ( 0.8 or 1.6 microsecond ) is below the amplitude of P3 by 6 dB. The Mode S transponder will reply this interrogation in the normal F1/F2 ATCRBS format. HAM US/F-4 SaR Dec 2005 Page 300 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 151 HAM US/F-4 SaR Dec 2005 All Call Interrogation Page 301 Part-66 Mode S interrogation The Mode S interrogation makes use of Differential Phase−Shift Keying (DPSK) modulation for the purpose of transmitting the high baud rate, 4 Million bits per second (4Mbps), data burst. Two general formats are possible: S short format with 56 bits of DPSK data or S the long format with 112 bits of DPSK data The signal comprises a P1, P2 preamble followed by a long pulse called P6. The P1, P2 pair preceding P6, suppresses replies from ATCRBS transponders to avoid synchronous garble caused by random triggering of ATCRBS transponders by the Mode S interrogation. A series of ’chips’ containing the information whithin P6 starts 0.5 microseconds after the sync phase reversal. A chip is an unmodulated interval of 0.25 microseconds duration preceded by possible phase reversals. The last chip is followed by a 0.5 microsecond guard interval which prevents the trailing edge of P6 from interfering with the demodulation process. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 302 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 152 HAM US/F-4 SaR Dec 2005 Mode S Interrogation Page 303 Part-66 Differential Phase Shift Keying (DPSK) DPSK−modulation is used to transmit bit coded information in pulse P6 of the uplink format. In the transmitter, a phaseshift of 180_ is generated. A 180_ phaseshift means logic ”1“ no phaseshift means logic ”O”. DPSK−demodulation is done in the receiver by delaying the previous signal (chip) which is 0.25 microseconds long by 0.25 microseconds, and comparing it with the actual signal received. If a phaseshift has occured, summing together delayed and actual signal will cause a 0 Volt output, representing a logic ’1’. Vice versa, if no phaseshift has occured, summing together delayed and actual signal will cause a double amplitude output, representing a logic ”0”. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 304 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 153 HAM US/F-4 SaR Dec 2005 Differential Phase Shift Keying (DPSK) Page 305 Part-66 Mode S Side Lobe Suppression Transmit sidelobe suppression is accomplished by the transmission of a control pulse (P5) on an SLS control pattern. If the control pulse amplitude received by the transponder exceeds the amplitude of the interrogation, the sync phase reversal will be obscured and the interrogation will be rejected. The P5 pulse must be used with the Mode S−only AllCall interrogation, to prevent unwanted replies from aircraft in the sidelobes. With discrete address interrogations, transmit SLS is not required to prevent sidelobe replies, as in general, an aircraft will be interrogated only when in the mainbeam of the interrogator antenna. However, transmit SLS on discretely addressed interrogations minimizes the probability of an aircraft erroneous accepting an interrogation directed to annother. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 306 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 154 HAM US/F-4 SaR Dec 2005 Mode S Side Lobe Suppression Page 307 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION ATC TRANSPONDER Part-66 Uplink Formats UF The Uplink Format (UF) consists of 25 Mode S interrogation formats (UF = 0 through UF = 24, as defined by RTCA/DO−181). The UF data is DPSK modulated and contained only in pulse P6 of Mode S interrogations. Presently only seven of the uplink formats (UF= 0,4,5,11,16,20 and 21) are defined and processed by the ATC mode S transponder. The first field (U #) in all uplink formats is the format number. It is coded in binary form. The last 24 bits (AP) of each uplink format represent the unique 24−bit address plus ovelaid parity bits. Abbreviations used in table Figure 155 : RL: (reply length): commands a reply in DF=O if zero, and a reply in DF=16 if one. AQ: (acquisition special): identifies UF=O and UF= 16 as acquisition transmis− sions and is repeated as received by the transponder in DF=0 and DF=16. MU: (message Comm−U): this 56−bit (33−88) uplink field contains information used in air−to−air exchanges and is part of the long special surveillance interrogation. P: (PAD−bits) PC: (protocol): contains operating commands to the transponder and is part of the surveillance and Comm−A interrogations. The codes are: S 0 = no changes in transponder state S 1 = non−selective All−Call lockout S 2 = not assigned S 3 = not assigned S 4 = cancel B S 5 = cancel C S 6 = cancel D S 7 = not assigned. RR:(reply request):length and content of the reply requested by the interrogations. S RR code 0 − 15 short reply S RR code 16 long reply, air−initiated Comm B S RR code 17 long reply, extended capability S RR code 18 long reply, flight ID S RR code 19 − 31 long reply. HAM US/F-4 SaR Dec 2005 DI: (designator identification): identifies the coding contained in the SD field. The codes are: S 0 = SD contains ILS, bfts 21−32 are not assigned S 1 = SD contains multisite information S 2 − 6 = not assigned S 7 = SD contains extended data readout request. SD: (special designator): contains control codes affecting transponder protocol. The content of this field is specffied by the DI field. MA: (message Comm A): contains messages directed to the aircraft. PR: (probabillity and reply): contains commands to the transponder that specify the reply probability to the Mode S−only All−Call interrogation (UF=11) that contains the PR. A command to disregard any lockout state can also be given. The assigned codes are as follows: S 0 = reply with probability =1 S 1 = reply with probability =1/2 S 2 = reply with probability =1/4 S 3 = reply with probability =1/8 S 4 = reply with probability =1/16 S 5 − 7 = do not reply S 8 = disregard lockout,reply with probability =1 S 9 = disregard lockout,reply with probability =1/2 S 10 = disregard lockout,reply with probability =1/4 S 11 = disregard lock out,reply with probability =1/8 S 12 = disregard lockout,reply with probability =1/16 S 13 − 15 = do not reply. Page 308 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION ATC TRANSPONDER Part-66 24 11 RC:2 NC:4 MC:80 Figure 155 HAM US/F-4 SaR Dec 2005 AP:24 Uplink Formats Page 309 Part-66 ATCRBS REPLIES The ground station assumes that the reply signal it receives after an interrogation signal is a valid response. The received reply signals are at a frequency of 1090 MHz. The ATCRBS reply format uses 13 serial pulses to convey aircraft identification in response to Mode A interrogations or aircraft altitude reporting in response to Mode C interrogations. The coded pulses are arranged between two framing pulses located at the beginning and the end of the data. The framing pulses F1 and F2 are spaced 20.3 microseconds apart. The encoding of the reply is done by means of the presence (1) or absence (0) of the 13 reply pulses (12 pulses plus X pulse). The code is delivered to the transponder from S switch settings on the control panel or S altitude settings coming from the Air Data Computers (ADC’s) in the aircraft. Only framing pulses will be transmitted, if code 0000 is selected for Mode A, or in Mode C if no altitude is available, or altitude reporting is switched to off. In addition to information pulses and framing pulses, a Special Position ldentification pulse (SPI) may be included in the reply by pushing the IDENT pushbutton, when requested by air trafic control. The SPI pulse follows the last framing pulse by 4.35 microseconds and is transmitted with each response that occurs during the 15 to 30 seconds period after the IDENT pushbutton has been actuated. The SPI pulse causes an enhanced pattern on the air traffic controller’s scope. Altitude encoding in replies Mode C replies from ATCRBS transponders encode the aircraft altitude in a gray code format that is also known as the Gilham code. The Gilham code is a 11 pulse format, that uses the presence of pulses to indicate increments in the altitude. In order to display negative altitude values down to −1000 feet, the 0 value starts at −1000 feet, instead of 0 feet. Refer to the table for examples of Gilham encoded altitudes. Mode S replies send the altitude information in the AC fields of formats 0, 4, 16, and 20 ( Refer to Figure 158 ) The 13 bit AC field expresses the encoded altitude in the sequence of C1, A1, C2, A2, C4, A4, M, B1, D1, B2, D2, B4, and D4. The M bit allows for the possible future use of encoding altitude in metric units. Zero is transmitted in each of the 13 bits if the altitude information is not available. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 310 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 156 HAM US/F-4 SaR Dec 2005 ATCRBS Replies Page 311 Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Part-66 MODE S REPLIES The Mode S reply is a PPM scheme. The format differs from the ATCRBS format and comprises four preamble pulses, followed by the reply data pulses (56 or 112 bits). In PPM coding, 0.5 microsecond pulses are placed within 1.0 microsecond windows, starting 8 microseconds after the leading edge of the first preamble pulse. If the positive pulse is located within the first half of the 1.0 microsecond window, the bit is understood to be a ’1’. If the positive pulse is located within the last half of the window, the bit is understood to be a ’0’. The downlink reply provides for 25 format possibilities. The 25 Mode S transponder transmission downlink formats DF = 0 through DF 24 are defined in RTCA/ DO−181. The short reply format 56 bits is used for the purpose of surveillance, squitter, and TCAS acquisition. The long format interrogations provide the same surveillance data as their short format counterparts but include provisions for 56 bits of non− surveillance type messages. For Training Purposes Only Mode S all call reply ( DF11 ) A Mode S all call reply is a response of a Mode S all call interrogation ( P4 duration 1.6 microseconds ) and consists of a four−bit preamble followed by a data block. The all call reply will be the Downlink Format 11. The data block contains 56 position modulated pulses. Mode S reply ( DF 0 - DF 24 ) A mode S reply occurs after a mode S interrogation or a mode S all call interrogation (P4 duration 1.6 microseconds). A mode S reply consists of a four−bit preamble followed by a data block. The data block contains 56 (short) or 112 (long) position modulated pulses. HAM US/F-4 SaR Dec 2005 Page 312 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 157 HAM US/F-4 SaR Dec 2005 Mode S Replies Page 313 Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Part-66 Mode S message content The main function of Mode S is surveillance. To accomplish this function, the Mode S transponder uses the 56−bit transmissions (in each direction). In the 56−bit transmissions, the aircraft reports its altitude or ATCRBS 4096 code, and the flight status (airborne, on−ground, alert, Special Position Identification (SPI), etc.). The discrete addressing and digital encoding of Mode S transmissions permit their use as a digital data link. The interrogation and reply formats of the Mode S system contain sufficient coding space to permit the transmission of data. These data transmissions may be used for air traffic control purposes, air−to−air data interchange for collision avoidance, or to provide flight advisory services such as weather reports, or Automated Terminal Information System (ATIS). Longer messages are transmitted using the Extended Length Message (ELM) capability. The ELM is capable of transmitting up to sixteen 80−bit message segments, either ground−to−air or air−to−ground. The table gives the definition of uplink format message fields: For Training Purposes Only Downlink Format Only seven downlink formats are defined and processed ( DF=0, 4, 5,11, 16, 20, and 21 ). The first field ( D # ) in all downlink formats is the format number. All discrete Mode S interrogations and replies (except the all−call reply) contain the 24−bit discrete address ( AP ) of the Mode S transponder upon which 24 error detection parity check bits are overlaid. In the all−call reply, the 24 parity check bits are overlaid on the Mode S interrogation address ( PI ) and the transponders discrete address is included in the text of the reply ( AA ). Abbreviations used in Figure 158 : VS: vertical status: is’0’ when airborne, is ’1’ on ground SL: sensitivity level: TCAS operating levels that determine the area of protected volume around a TCAS−equipped aircraft, as stated in DO−185 RI: reply information: reports airspeed capability and type of reply to the interrogating aircraft. The coding is as follows: S 0−7 = codes indicate that this is the reply to an air−to−air non−acquisition interrogation S 8−15 = codes indicate that this is an acquisition reply S 8 = no maximum airspeed available HAM US/F-4 SaR Dec 2005 S 9 = airspeed is less than or equal to 75 knots S 10 = airspeed is greater than 75 and less than 150 knots S 11 = airspeed is greater than 150 and less than 300 knots S 12 = airspeed is greater than 300 and less than 600 knots S 13 = airspeed is greater than 600 and less than 1200 knots S 14 = airspeed is greater than 1200 knots S 15 = not assigned AC: altitude code: Contains altitude information. Zero is transmitted in all 13 bits if altitude information is not available. Metric altitude is contained in this field if bit 26 is set to ’1’. MV: message, Comm−V: contains information used in air−to−ground exchanges UM: utility message: Contains transponder status readouts. MB: message Comm−B:Contains messages to be transmitted to the interrogator and is part of Comm−B replies DF=20 and 21. BDS: Comm−B definition subfield: Contained within MB. Defines the content of the MB message field of which it is a part FS: flight status: reports the flight status of the aircraft. DR: downlink request: used to request extraction of downlink messages from transponder by the interrogator. ID: identification:contains the 4096 identification code reporting the number set at the control unit CA: transponder capability: reports transponder capability. Codes are: S 0 = no communications capability (surveillance only). Transponder accepts UF= 0, 4, 5, 11 Transponder transmits DF = 0, 4, 5, 11 S 1,2,3 = indicates the specific data link capabilities of the overall airborne installation S 4−7 = not assigned AA: address announced: contains the aircraft address PI: parity/interrogator identity: contains the parity overlaid on the interroga− tor’s identity code Page 314 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION ATC TRANSPONDER Part-66 24 11 -1- KE:1 ND:4 MC:80 Figure 158 HAM US/F-4 SaR Dec 2005 AP:24 Downlink Formats Page 315 Part-66 System Architecture The ATC comprises two independent systems. Each system consists of: −one transponder, −two antennas, −one ATC/TCAS control unit, common to the two systems. In addition, a Centralized Fault−Display Interface−Unit (CFDIU) enables access to the maintenance part of the ATC system through one Multipurpose Control and Display Unit (MCDU). The system is provided with four antennas : S two antennas are located at the upper part of the fuselage S two antennas are located at the lower part of the fuselage. According to the aircraft configuration, the transponder selects the antennas (lower or upper) which receive the best transmission signal from the ground ATC secondary radar. Warning The warning related to the ATC is only the FAULT or FAIL indicator light of the ATC control unit : S when the system is faulty, the amber FAULT or FAIL indicator light comes on. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 316 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 159 HAM US/F-4 SaR Dec 2005 ATC System Schematic Page 317 Part-66 ATC TYPICAL LOCATION For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 160 HAM US/F-4 SaR Dec 2005 ATC Locations Page 318 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 161 HAM US/F-4 SaR Dec 2005 ATC Locations Page 319 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Part-66 COMPONENT DESCRIPTION Because there are a lot of different layouts for the components, the following description will only be one example. ATC Transponder S External description The face of the transponder is fitted with a handle, two attaching parts, a TEST pushbutton switch and ten LEDs. The name, color and function of the ten LEDs are as follows: − TPR (green) indicates that no faults are detected during the reply − TPR (red) indicates fault of ATC function − ANT TOP (red) indicates that the top antennas are incorrectly connected or are failed − ANT BOT (red) indicates that the bottom antennas are incorrectly connected or are failed − ALT (red) indicates that the altitude is not in correct format − DATA IN (red) indicates that the data input is not in correct format − TCAS (red) indicates that the data input from the TCAS is abnormal − MAINTENANCE (red) indicates that the maintenance data are abnormal − RESERVED (red) are reserved for future use. The back of the transponder contains an ARINC 600, shell size connector to provide electrical connections to the aircraft wiring via mount. Contact grouping is as follows: − Top contact set − ATE interface − Center contact set − System interconnections − Bottom contact set − Power supply and bonding. HAM US/F-4 SaR Dec 2005 S Principle of operation Each ground interrogator transmits its interrogations at the frequency of 1030 MHz in the form of a series of two pulses. Depending on the pulse intervals and numbers, they define three different interrogation modes. After receiving these pulses, the transponder identifies and decodes the interrogations. Depending on the detected interrogation mode, the transponder transmits either the identification of the aircraft or its barometric altitude or flight data. To perform these functions, the transponder is associated with a control unit defining the aircraft identification code, the Air Data/Inertial Reference Units (ADIRUs) the Flight Management and Guidance Computers (FMGCs) and the Traffic Collision Avoidance System (TCAS). The transmission of the replies to the ground takes place on a carrier frequency of 1090 MHz. If the interrogation is sent by the side lobe of the radar a characteristic signal is sent allowing the transponder to disregard the interrogation. In addition to its specific transponder functions, it enables communication between the TCAS and a detected aircraft if equipped with the TCAS. Three classes of interrogations are transmitted by the ground interrogator, they are: −Mode A or C interrogations −Mode S all call interrogations −Mode S interrogations. Page 320 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 162 HAM US/F-4 SaR Dec 2005 ATC Transponder Page 321 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Part-66 ATC / TCAS Control Unit The ATC control panel allows the pilot to select the proper mode of operation, to select codes for reply and to test and monitor system operation. The processor−based unit conforms to the ARINC 718 specification. The front panel is integrally lit by clear 5V lamps and features : S a liquid crystal display window in which the identification code is displayed. Upon energization of the aircraft or modification of system selection, the last ATC code selected is displayed before the new code selection. This code comes into view on the display window of the control unit. S an ATC FAIL indicator light which indicates a transponder failure S a IDENT pushbutton switch. When you push the IDENT pushbutton switch (ON configuration), this generates an additional identification pulse in the reply signal. S an ALT RPTG/ON/OFF switch which controls the transmission of altitude data. When you set the ALT RPTG/ON/OFF switch to ON, the altitude information transmission of the selected transponder is enabled. In normal configuration, each ATC receives the altitude information from its corresponding Air Data/Inertial Reference Unit (ATC1 from ADIRU1, ATC2 from ADIRU2). With failure of the ADIRU corresponding to the serviceable transponder, the pilot can select the altitude information from the ADIRU 3. This selection is through the AIR DATA selector switchl. S a keyboard to select the four−digit transmit code (from 0000 to 7777) S a 1/2 selector switch which enables the operation of the ATC system 1 or 2 S a STBY/AUTO/ON selector switch which enables the ATC operating modes − STBY: Selected ATC Transponder is electrically supplied, but not operating. (The answers of the two transponders are inhibited ) − AUTO: The selected ATC Transponder answers are automatically inhibited when the aircraft is on the ground. On ground only DF11 as Squitters are transmitted. − ON: Selected ATC Transponder answers are valid when the aircraft is in flight or on the ground. S a TCAS mode selector switch (STBY/TA/TA−RA) which enables the ATC/ TCAS operating modes (Ref. TCAS for more details). S a THRT/ALL/ABV/BLW selector switch which controls the type of TCAS-traffic to be displayed HAM US/F-4 SaR Dec 2005 Each transponder can substitute the other in case of failure. The selected tansponder reads the switch selections and : S drives the LCD display S formats the ARINC control word. In case of transponder failure, two independent circuits control : S the failure indicator light S the failure mode of the processor. With failure of the ATC/TCAS control unit, the transponder continues to transmit the last displayed code. No new code selection is possible. Page 322 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 163 HAM US/F-4 SaR Dec 2005 ATC / TCAS Cont. PNL Page 323 Part-66 ATC Antenna The L−band antenna is a short stub, all aluminum blade type which is completely sealed to prevent failure from moisture incursion. It is vertically polarized, has an impedance of 50 ohms, a Voltage Standing Wave Ratio ( VSWR ) of 1,5 : 1 and operates in the 960 MHz to 1220 MHz frequency band. Lightning protection is provided to prevent damage to the antenna and the transponder. Four antennas must be provided on the aircraft. The desire for improved upper hemisphere coverage usually leads to a choice of upper and lower mounting locations for the transponder antennas on the aircraft. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER HAM US/F-4 SaR Dec 2005 Page 324 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION ATC TRANSPONDER Figure 164 HAM US/F-4 SaR Dec 2005 ATC Antenna Page 325 Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Part-66 ATC MODE S TRANSPONDER SELFTEST A built−in test function allows the transponder to monitor its own circuits, the upper and lower transponder antennas, the air data source output, and the TCAS/transponder control panel. Monitoring is continuous during operation. Selftesting may be initiated manually and depends on the system configuration. The transponder can manually be tested by: S engaging the test switch on the TCAS/ATC control panel or S depressing the transponder front panel TEST switch. Front panel indicators will light randomly for a brief period following initiation of self−test. As testing continues indicators will turn on and off momentarily, then the indicator(s) associated with failed LRUs will light just before all indcators turn off. The following identifies the front panel indicators. For Training Purposes Only FUNCTION / RELATED LRU INDICATOR −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− TPR PASS Successful BITE self−test of transponder TPR FAIL Failed BITE self−test of transponder UPPER ANT Faulty impedancy in upper antenna LOWER ANT Faulty impedancy in lower antenna ALT Faulty data output or format from selected ADC CTL Faulty data output or format from TCAS / ATC control panel S by using the maintenance tests with the centralized fault−display system (CFDS). The primary components of the CFDS are : S the Multipurpose Control and Display Unit (MCDU) S the Centralized Fault Display Interface Unit (CFDIU) S the Built−In Test Equipment (BITE) internal to the ATC transponder. It is possible to perform these tests on the ground only. To get the ATC 1(2) page on the MCDU: − select of the SYSTEM REPORT/TEST item on the CFDS menu − then, push the line key adjacent to the ATC 1(2) indication on the NAV menu. The test sequence starts when you push the line key adjacent to the TEST indication on the ATC 1(2) page. After the INTERFACE TEST selection, the message : PLEASE CHECK IF THE CONTROL PANEL SWITCH IS ON SYSTEM (1 OR 2) AND PRESS TEST KEY is displayed on the MCDU. HAM US/F-4 SaR Dec 2005 Page 326 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION ATC TRANSPONDER Figure 165 HAM US/F-4 SaR Dec 2005 ATC CFDS Menu Page 327 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 TCAS GENERAL Presentation The TCAS II (Traffic Collision Avoidance System) is a system whose function is S to detect and S to display For Training Purposes Only aircraft in the immediate vicinity and S to provide the flight crew with − indications and − aural messages to avoid these intruders by changing the flight path in the vertical plane only. The TCAS periodically interrogates their transponders, computes their trajectories and constantly determines their potential threat. Their acquisition is achieved by means of two transmit/receive antennas, one located on the underside of the fuselage and the other on the top. The system can establish individualized communications with each aircraft through ATC/Mode S transponders, thus permitting operation in dense traffic areas while avoiding an overload of radio transmissions that would result from a general all−intruder response. The TCAS II system is designed to provide the air traffic control system: it usually operates independently but may be also controlled from ground stations. The system maintains surveillance within a sphere determined by the transmit power and receiver sensitivity of the TCAS computer. The area in which a threat is imminent depends on the speed and path of the own A/C and the threat A/C. The TCAS detection capability covers an area of S about 30 NM in range and S plus or minus 9000 ft ( 9900 ft ) in altitude Display range normally is authorized up to plus or minus 1200 ft in altitude only. HAM US/F-4 SaR Dec 2005 Page 328 Part-66 ( 9900 ft ) For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS ( 9900 ft ) Figure 166 HAM US/F-4 SaR Dec 2005 TCAS-Surveillance Envelope Page 329 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 INTRODUCTION Principle The traffic alert and collision avoidance system (TCAS) helps the flight crew maintain safe air traffic separation from other ATC transponder equipped airplanes. TCAS is an airborne system and operates independently of the ground−based ATC system. TCAS sends interrogation signals to nearby airplanes. These airplanes, which are equipped with a ATCRBS ( mode A/C ) or a mode S transponder, respond to these interrogations. TCAS uses these response signals to calculate: S range, S relative bearing, and S altitude of the responding airplane. If a responding airplane does not report altitude, TCAS cannot calculate the altitude of that airplane. Airplanes tracked by TCAS are called targets. Using the information from the response signals and altitude of own airplane, TCAS calculates the closure speed between own airplane and the target. TCAS then calculates how close the target will be to own airplane at the closest point of approach (CPA). For Training Purposes Only There is an area defined as TAU within the surveillance arc, which represents the minimum time the flight crew needs to discern a collision threat and take evasive action. Targets are classified as one of these four types depending on the separation at CPA and the time it will take until CPA occurs: S Other traffic S Proximate traffic S Intruders TA S Threats RA Each type of target has a different symbol on the display. HAM US/F-4 SaR Dec 2005 Page 330 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 PROXIMATE { 20/40 sec 35/45 sec For Training Purposes Only 6 NM Maxi range of the display 30 NM Figure 167 HAM US/F-4 SaR Dec 2005 Range Closure Rate Time to CPA TCAS Target Classification Page 331 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION TCAS Part-66 TCAS Encounter If the separation at CPA is within certain limits, TCAS provides advisory messages to the flight crew. TCAS provides two levels of advisories to the flight crew, S traffic advisory (TA) and S resolution advisory (RA). The type of advisory is determined by a combination of altitude, the time to CPA, and the separation at CPA. The TA shows for relatively longer times to CPA and relatively larger separation at CPA and is for intruder targets. The TA shows the range, bearing, and relative altitude (if relative altitude is known) of the intruder target and gives aural TRAFFIC-TRAFFIC announcements. The RA shows for relatively shorter times to CPA and relatively smaller separation at CPA and is for threat targets. The RA also gives visual and aural commands to the flight crew to make sure there is safe vertical separation from the threat target. TCAS also communicates with other airplanes that have TCAS to coordinate the flight movement to prevent a collision. Display Visual indications are presented on the Electronic Flight Instrument System (EFIS) or dedicated displays. The Navigation Display (ND) is used to indicate the situation in the nearby traffic area : a symbol is displayed for each intruder on the image. The avoidance maneuver indications, if any, are displayed on the vertical speed scale of the PFD by means of a band of colored sectors showing the vertical speed value to be adopted in order to avoid any risk of collision. Advisories Visual and aural advisories are supplied by the TCAS computer whenever assessment of the relative position of two aircraft reveals a potential collision hazard. The Traffic Advisories (TA) indicate the position of nearby aircraft which are or may become a threat. Their display alerts the flight crew to the presence of intruders and facilitates their visual acquisition. The Resolution Advisories (RA) may be divided into two categories: S Corrective Advisories that instruct the pilot to deviate from current vertical rate S Preventive Advisories that instruct the pilot to avoid certain maneuvers. HAM US/F-4 SaR Dec 2005 Page 332 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 168 HAM US/F-4 SaR Dec 2005 TCAS encounter Page 333 Part-66 TCAS Coordination The avoidance maneuvers initiated by the TCAS could create a conflict situation if directed at another TCAS−equipped aircraft as this aircraft may also take similar evasive action, resulting in an unchanged situation. To avoid this situation, a communication link between the two aircraft is necessary, exchanging coordination messages. The first aircraft to detect the other one initiates the communication procedure, indicates the maneuvers it intends to perform and communicates orders to the other aircraft requesting it to maintain its trajectory. This necessarily involves the use of Mode S transponders, the only equipment of this type possessing the LINK function required for data exchange. The Mode S transponders provide the capability to transmit a unique address (24 bits) assigned to each aircraft, permitting them to reply individually to other TCAS− equipped aircraft. It can respond to ground station interrogations in Mode A and Mode C and also in Mode S if the stations are suitably equipped. With respect to aircraft equipped with Mode A transponders only, the TCAS cannot generate resolution advisories, as these transponders do not communicate aircraft altitude. The aircraft are however displayed on the ND, enabling their location in range and bearing, but with no altitude indications. The ground stations can modify the TCAS operating mode via the transponder link so as to inhibit resolution advisories in certain conditions. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS HAM US/F-4 SaR Dec 2005 Page 334 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 169 HAM US/F-4 SaR Dec 2005 TCAS Coordination Page 335 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 GENERAL DESCRIPTION These are the TCAS components: S TCAS computer S ATC/TCAS control panel. S TCAS directional antennas (2) For Training Purposes Only TCAS interfaces with these other system components: S ATC transponders (2) S Landing gear lever S Ground Sensing Relay S Display electronic units (DEUs) or Display Management Computer (DMC) S Remote electronics unit (REU) or Audio Management Unit (AMU) S Radio altimeters (2) S Ground proximity warning computer (GPWC) S Weather radar S Air data inertial reference unit (ADIRU) S Suppression coax tees S FDAU or FDIU. HAM US/F-4 SaR Dec 2005 Page 336 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 TOP ATC ANTENNA TOP TCAS DIRECTIONAL ANTENNA BOTTOM ATC ANTENNA TO DISPLAYS DEU or DMC REU or AMU ATC TRANSPONDER (2) RADIO ALTIMETER TA/RA TA ABV OFF BLW T C A S 1 2 3 ATC1 TEST IDENT 4 7 5 0 6 XPNDR CLR 1 2 AUTO STBY ON A T C GPWC FAULT ADIRU For Training Purposes Only ATC/TCAS CONTROL PANEL TCAS COMPUTER LANDING GEAR LEVER GROUND SENSING RELAY FDAU or FDIU BOTTOM TCAS DIRECTIONAL ANTENNA Figure 170 HAM US/F-4 SaR Dec 2005 SUPPRESSION COAX TEES TCAS General System Schematic Page 337 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 TCAS- PWR, ANTennas, ANALOG-, DISCRETE INTERFACES General The TCAS computer has analog and discrete interfaces with these components: S Top and bottom TCAS directional antennas S Landing gear lever switch S Proximity Switch Electronics Unit (PSEU) S Ground proximity warning computer (GPWC) S Weather radar S DME/ATC/TCAS suppression coax tees S Display electronic unit (DEU) 1 and 2 S Remote electronics unit (REU). Power The TCAS computer gets 115v ac from AC transfer bus−1 through the TCAS circuit breaker on the P18 circuit breaker panel. Antennas There are two TCAS directional antennas. The TCAS directional antennas receive traffic airplane reply signals. They also transmit the TCAS interrogation signals. For Training Purposes Only Landing Gear Lever Switch The discrete from the landing gear switch tells the TCAS computer that the landing gear is down. When the TCAS computer gets this discrete, the TCAS computer makes the bottom directional antenna become an omnidirectional antenna. PSEU The discrete from the PSEU supplies in−air or on−ground status to the TCAS computer. The air/ground discrete inhibits TCAS operation on the ground and inhibits tests when in the air. The air/ground discrete also controls flight leg increments in the TCAS nonvolatile memory. GPWC − Advisory Inhibit Discretes The GPWC sends three discretes to the TCAS computer. These discretes inhibit TCAS aural and visual warnings during all GPWS modes except mode 6 (altitude callouts). HAM US/F-4 SaR Dec 2005 Weather Radar The TCAS computer gets one discrete from the weather radar. This discrete inhibits all TCAS aural warnings and changes RAs to TAs when the weather radar makes a predictive windshear warning. Suppression Input/Output The TCAS computer gets a suppression pulse when an ATC transponder or DME interrogator transmits. When the TCAS computer transmits, it sends a suppression pulse to the ATC transponders and the DME interrogators. DEU − Display Status A discrete from either DEU goes to the TCAS computer when the DEU loses the ability to show TCAS displays. When the TCAS computer gets this discrete, the TCAS computer does not do these functions: S Send TCAS display outputs to the DEU S Send TCAS aurals to the REU S Transmit coordination data to traffic airplanes with TCAS. REU − TCAS Aurals − Voice Outputs The TCAS computer sends resolution advisory (RA) and traffic advisory (TA) aural signals to the remote electronic unit (REU). The REU amplifies the RA and TA aurals. Then it sends them to the flight interphone speakers and headsets to alert the flight crew. Program Pins Program pins control the configuration of the TCAS computer. These are the functions that the program pins on the TCAS computer permit: S Shows a maximum of 8 targets S Goes to the standby mode when the airplane is on the ground. S Does not show airplanes that are on the ground when own airplane is below 1750 feet AGL S Inhibits self−test in the air S Controls the audio level of the voice outputs S Sets the airplane altitude limit of 48,000 feet so TCAS does not command a climb or increase climb above this altitude. Page 338 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 115V AC XFR BUS 1 TCAS ATC 1 TRANSPONDER P18 CIRCUIT BREAKER PANEL RF TRANSMIT AND RECEIVE SUPPRESSION COAX TEE DME 1 TOP TCAS DIRECTIONAL ANT COAX TEE ATC 2 TRANSPONDER RF TRANSMIT AND RECEIVE BOTTOM TCAS DIRECTIONAL ANT GEAR UP COAX TEE RA STATUS LANDING GEAR LEVER SW DEU 2 AIR AIR/GND STATUS GROUND For Training Purposes Only DEU 1 LANDING GEAR STATUS GEAR DOWN TCAS RA AND TA AURALS REU PSEU PROGRAM PINS WINDSHEAR WARNING ALERT GPWC TCAS COMPUTER Figure 171 Dec 2005 MAX INTRUDER DISPLAY 8 STANDBY ON GROUND ON GRD INTRUDER DISABLE SELF TEST INHIBIT AUDIO LEVEL CONTROL PROGRAM COMMON 32,000 FT 16,000 FT PROGRAM COMMON WARNING HAM US/F-4 SaR DME 2 RA STATUS TCAS − PWR, A/ D INTERFACES AND PPs Page 339 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 DIGITAL INTERFACES General The TCAS computer has digital interfaces with these components: S ATC 1 transponder 1/2 S Radio altimeter 1/2 transceiver S Left air data inertial reference unit (ADIRU) S Display electronic unit 1/2 (DEU 1/2) S Flight data acquisition unit (FDAU). TCAS Inputs from ADIRU The left ADIRU supplies these inputs to the TCAS computer: S Airplane roll attitude S Airplane pitch attitude S Airplane heading. TCAS Outputs to DEUs The TCAS computer supplies resolution advisory (RA) and traffic advisory (TA) data to the DEUs. This includes all traffic data for TCAS displays. TCAS Outputs to the FDAU The FDAU receives the same TCAS data that goes to the DEUs. For Training Purposes Only ATC Transponder − Control and Coordination Data When you select an ATC transponder, it sends this ATC control panel data to the TCAS computer: S The TCAS mode selection (TA only or TA/RA) S Control of the altitude limits for the TCAS display that shows on the navigation display (ND). The TCAS computer uses this data from the ATC transponder to calculate the climb or descent to prevent a collision: S 24−bit airplane address S Barometric altitude S Maximum true airspeed. The TCAS computer sends this to the ATC transponder: S General TCAS operational status S Mode S coordination data. Radio Altimeter Inputs The TCAS computer gets radio altitude from radio altimeter 1 and 2 transceivers although only one input is necessary for TCAS operation. The data is used by the TCAS computer to calculate some sensitivity levels. At approximately 1700 feet, the TCAS computer uses radio altitude together with barometric altitude to determine intruders that are on the ground and therefore no threat to the TCAS airplane. At 1000 feet radio altitude, the TCAS computer inhibits resolution advisories and TA ONLY will show on the NDs. HAM US/F-4 SaR Dec 2005 Page 340 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 For Training Purposes Only Figure 172 HAM US/F-4 SaR Dec 2005 TCAS − DIGITAL INTERFACES Page 341 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION TCAS Part-66 COMPONENT DESCRIPTION TCAS COMPUTER Purpose The TCAS computer is the main component of the TCAS. It controls these functions: S Surveillance S Tracking S Advisory S Air−to−air maneuver coordination. The TCAS computer sends signals which tell the flight crew to make one of these maneuvers: S Keep the current flight plan S Make flight maneuvers to prevent a possible collision with other airplanes. Physical Description The TCAS computer is a 6 MCU size unit. It weighs 28 lbs (11.3 kg). Functional Description The TCAS computer transmits 1030 MHz pulse−coded interrogation signals. It receives 1090 MHz pulse−coded reply signals from intruder airplanes with an ATC transponder. Front Panel LCD Indications Two alphanumeric LCDs show fail conditions of the TCAS computer and systems that have interfaces with the TCAS computer. Front Panel Self−Test You push the test switch on the front panel to start a test of the TCAS. The LCD indications come on to show system status. Front Panel Connectors There are two connectors on the front panel. One is for automated test equipment and one is to load software . The TCAS computer is supplied with the software fully loaded. Nevertheless, the plug is installed for the loading of the operational program and I/O configuration data into the TCAS computer via 2 ARINC 429 low speed buses, by means of a data loader. HAM US/F-4 SaR Dec 2005 Page 342 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 173 HAM US/F-4 SaR Dec 2005 TCAS Processor Page 343 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 TCAS − ATC/TCAS CONTROL PANEL General The TCAS is a cooperative system whose operating mode is very close to the ATC Mode S transponder associated to it. The main controls are thus grouped on the ATC/TCAS control unit and the traffic and conflict resolution information is presented on the EFIS displays. The manual operating modes of the TCAS are selected via the ATC/TCAS control unit. TCAS modes of operation The TCAS mode of operation is selected by means of two selector switches: S STBY, TA and TA/RA − STBY mode In this mode, TCAS intruders (proximate or/and other) are displayed if a TA or a RA is already displayed. − ALL mode This selection enables display of all intruders without any conditions (TCAS intruders are displayed when detected). − ABV and BLW modes This selection controls the above and below vertical altitude for traffic advisory: In the Standby Mode, the advisory generation and surveillance functions are inhibited. No TCAS information can be displayed on the PFDs and NDs. ABV: altitude range is set to 9900 ft above the aircraft and 2700 ft below The aircraft symbol and the range ring remain on the ND and vertical speed information is not displayed on the PFD. BLW: altitude range is set to 9900 ft below the aircraft and 2700 ft above. Operation of the ATC section is described under ATC subject. The green TCAS STBY message is displayed in the memo section of the EWD. − TA mode For Training Purposes Only S THRT, ALL, ABV and BLW. − THRT mode Note: In this mode, intruders are displayed on the ND according to their position in the airspace. The RA type intruder symbols are converted into TA type symbols. The TCAS performs surveillance functions but does not generate any resolution advisories. The TA ONLY message is displayed in white on the NDs in the left corner of the TCAS message area. − TA/RA mode The TCAS performs all TA mode functions and also issues preventive or corrective resolution advisories, represented in the form of colored sectors along the vertical speed scale on the PFD. The sensitivity level is determined automatically in function of altitude. HAM US/F-4 SaR Dec 2005 Page 344 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 174 HAM US/F-4 SaR Dec 2005 TCAS / ATC Cont. PNL Page 345 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 TCAS Antennas General The TCAS uses a top and bottom directional antenna. The antennas are the same and interchangeable. The directional antenna is composed of four passive vertically−polarized elements. This high−strength composite antenna is provided with a flat base, fuselage mounting screws and four color−coded connectors used to coaxially connect the four antenna elements to the TCAS computer. The antenna is used to receive and provide directional information for 1090 MHz Mode S squitters, Mode S and Air Traffic Control Radar Beacon System (ATCRBS) replies. Proper phasing of the four antenna elements enables omni or directional transmission of 1030 MHz broadcast or coordination messages and ATCRBS or mode S interrogations. CAUTION: DO NOT PAINT THE RADIATION SURFACE OR THE BACKPLATE OF THE ANTENNA. PAINT DOES NOT PERMIT THE ANTENNA TO RADIATE 0R RECEIVE RF SIGNALS. CAUTION: TO PREVENT DAMAGE TO THE ANTENNA CABLES, DO NOT PULL ON THEM. How is bearing determined LEERER MERKERdisplays the TCAS transmit pattern for the area directly ahead of the aircraft (front quadrant). The TCAS II transmits interrogation signals to other aircraft in a manner similar to a ground based ATC station. The ground based ATC station uses a narrow beam, mechanically rotated antenna, while the TCAS II system uses a wider, electrically directed beam. The directionality is accomplished by inserting a different electrical phase delay into each of the rf signals applied to the elements of the antenna. LEERER MERKERshows the phase delay for the four possible directions and an omnidirectional pattern. After interrogating, the TCAS II system then waits to receive a reply (”listens”) from the direction the transmission was sent. The phase difference in the received signals between the two pairs of antenna elements is used to determine the replying intruder bearing. The antenna elements used in determining the phase difference are determined by the interrogation direction. For example, if the intruder is approaching in the front quadrant, then the TCAS II calculates the phase difference between elements 1 and 2 and also 1 and 4. The relationship between the two phase differences then corresponds to the intruder bearing. For Training Purposes Only Training Information Point An antenna connection includes the coax cable and an antenna element. The TCAS computer checks the resistance of each antenna connection at power—up. The TCAS computer reports an antenna fault when it detects that the resistance of the connection is out of range. If you do not connect the coax cable to the correct element, the TCAS computer reports an antenna fault. HAM US/F-4 SaR Dec 2005 Page 346 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 175 HAM US/F-4 SaR Dec 2005 TCAS ANTENNA Page 347 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION TCAS Part-66 OPERATION Intruder Detection The TCAS detects A/C equipped with Mode S transponders by listening for squitter transmissions. Mode S transponders announce their presence by transmitting squitter messages once every second. The TCAS also detects A/C, that do not reply to Mode S interrogations but do reply to Mode C interrogations. The TCAS must actively search for Mode C equipped intruder aircraft because Mode C transponders do not transmit squitter messages. Once the presence of a Mode C intruder is confirmed, it is tracked by the TCAS. The TCAS is capable of tracking up to a combined total of 30 Mode S and Mode C intruders. Tracking is performed by repetitive TCAS interrogations in Mode S and Mode C format. S Interrogation of aircraft equipped with Mode A or C transponders. With respect to aircraft equipped with Mode A or Mode C transponders, the TCAS is active and transmits Mode C only all−call interrogations (P1, P3 and P4 pulses). The code is similar to the one used by the Mode A and Mode C ground stations. The P4 pulse informs those Mode S transponders that this interrogation is not addressed to them. − Whisper-Shout The nominal time interval between two interrogations is one second. But, to limit radio−electric interference in dense traffic areas, each interrogation consists of a series of interrogations of increasing strength to reach more remote aircraft ( whisper−shout ) with 1 ms time periods inside the series. The first transmission consists of relatively low power P1, P3 and P4 pulses only. Therefore, only the nearest aircraft will receive and reply to these interrogations. Then an S1 pulse is also transmitted. This pulse is at a lower amplitude, causing the close−in aircraft to interpret this as a side lobe from the transmitting station, requiring no reply. The purpose of the whisper− shout sequence is to reduce the number of aircraft replying to any one interrogation, thus limiting interference. HAM US/F-4 SaR Dec 2005 Page 348 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 Mode C Only All-Call interrogation For Training Purposes Only Whisper-Shout Transmitter Sequence Figure 176 HAM US/F-4 SaR Dec 2005 TCAS Whisper-Shout Page 349 Part-66 S Replies of A/C equipped with Mode A or Mode C transponders. Aircraft equipped with Mode C transponders reply by transmitting their altitude, octal encoded in four digits ABCD, with a value of 100 ft for the LSB in the ATCRBS format. Aircraft equipped with Mode A transponders reply by transmitting their Mode A identification code. In this case, intruder presentation on the ND is limited to a display of its position in range and bearing. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS HAM US/F-4 SaR Dec 2005 Page 350 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 177 HAM US/F-4 SaR Dec 2005 Transponder Mode A/C Reply Format Page 351 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 S Interrogation of aircraft equipped with Mode S transponders. − Transmission coding The TCAS uses the Mode S function for certain identification of intruders as a 24−bit address is definitively assigned to each aircraft by air traffic control. The interrogation comprises three pulses : P1, P2 and P6. P2 level is equal to or greater than the P1 level, which is the no−reply condition for the aircraft equipped with Mode A or C transponders. Therefore, only Mode S transponders reply to the interrogation. The useful information is contained in P6 divided into 56 or 112 chips. A chip is an unmodulated interval of 0.25 microseconds, preceded by possible phase reversals. The message formats contain a number of bits permitting a more complete and diversified information exchange than in Mode C. There are two distinct message formats : −all Mode S interrogations (UPLINK format) are binary differential phase shift keying (DPSK) signals For Training Purposes Only −Mode S replies (DOWNLINK format) are formed by pulse position modulation (PPM) encoding the reply data. The Mode S reply is preceded by a preamble containing four pulses of specific duration and intervals intended to guarantee received message validity. Any messages whose preamble is not in complete conformity with the model are rejected by the TCAS. This information is encoded in PPM mode with, for each bit, a logic level one if the first half of the interval is at 1 and a logic level zero if it is at zero. − Squitters The Mode S transponder participates actively in its own detection by transmitting signals, at one second intervals, intended to inform nearby aircraft of its presence. This transmission, called squitter, consists of a Format DF = 11 message, containing the Mode S 24−bit address assigned to the aircraft, whereas all the bits of the message PI field at zero indicates a squitter. HAM US/F-4 SaR Dec 2005 Page 352 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 UF For Training Purposes Only DF Figure 178 HAM US/F-4 SaR Dec 2005 TCAS -Mode S-Interrogation and Reply Page 353 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 − Mode S communication messages The Mode S ATC communications system supporting personalized exchanges between ATC ground stations and Mode S transponder−equipped aircraft comprises : −a set of 25 standard messages for station−aircraft uplinks (Uplink format) identified UFxx −another set of 25 standard messages for downlinks (Downlink for mat) identified DFxx.The message may be long or short and contain either 56 or 112 bits. Each message consists of specific fields with bit combinations that have specific meanings. Two fields, however, have an identical definition for all messages : −the ”message−type” field consisting of bits 1 to 5 whose coding translates the format decimal value into binary −the address field containing the last 24 bits of the message (bits 33 to 56 or bits 89 to 112 depending on whether the message is short or long). These 24 bits contain the Mode S address of the transmitter. The TCAS uses this bidirectional link capability to communicate with other TCAS equipped aircraft to coordinate avoidance maneuvers. It may also dialog with ground stations : these stations have the possibility of monitoring and modifying its action. In the Mode S message set, the TCAS uses only six UF−type messages and seven DF−type messages. These messages are : For Training Purposes Only −UF0, UF4, UF5, UF16, UF20, UF21 −DF0, DF4, DF5, DF11, DF16, DF20, DF21. The Figure 179 give the list of messages used by the TCAS for communications with other aircraft and with ground stations. The two tables give the definition of the fields used in the messages. HAM US/F-4 SaR Dec 2005 |DESIGN | FIELD | INDICATION | |−−−−−−−|−−−−−−−−−−−−−−−−|−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | | AP |ADDRESS PARITY |Coded address with parity check | | AQ |ACQUISITION |Indicates if it is an interrogation message ; | | | |1 = interrogation | | DI |DESIGNATOR |Specifies type of information contained in | | |IDENT |SD field | | MA |MESSAGE |Used by ground station to transmit a TCAS | | |Comm−A |SLcommand to a TCAS−equipped aircraft | | MU |MESSAGE |Used by TCAS to transmit to other aircraft | | |Comm−U |RA coordination information (under fields | | | |UDS,MTB, CVC,VRC,CHC,HRC,HSB, VSB) | | PC |PROTOCOL |Operating commands to the transponder | | RL |REPLY LENGTH |Indicates if message is short (0) or long (1) | | RR |REPLY REQUEST |Length and content of reply information | | | |requested by the interrogator | | SD |SPECIAL |Contains control codes affecting the | | |DESIGNATOR |transponder protocol | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− Table 1: N of uplink format message fields |DESIGN | FIELD | INDICATION | |−−−−−− |−−−−−−−−−−−−−−−− |−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | AA |ADDR ANNOUNCED |Mode S address in the clear in 24 bits | | AC |ALTITUDE CODE |Information indicating aircraft altitude | | AP |ADDRESS PARITY |Coded address with parity check | | CA |CAPABILITY |Transponder capability | | DR |DOWNLINK |Requests extraction of downlink message by| | |REQUEST |the interrogator (existing RA) | | FS |FLIGHT STATUS |Flight status of the a/c: gnd,flight, alert,SPI | | ID |IDENT CODE |Contains the Mode A identification code | | MB |MESSAGE |Indicates Advisory content to the ground | | |Comm−B |station | | MV |MESSAGE |Contains ARA, RAC, VDS subfields used | | |Comm−V |for coordination | | RI |REPLY |INFO |Type of reply and airspeed capability | | SL |SENSITIVITY LEVEL |TCAS current sensitivity level | | UM |UTILITY MESSAGE |Transponder status readouts | | VS |VERTICAL STATUS |Aircraft status : 0 = airborne, 1 = ground | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− Table 2: N of downlink format message fields Page 354 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 179 HAM US/F-4 SaR Dec 2005 TCAS used UP- / Downlink Formats Page 355 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION TCAS Part-66 MEASUREMENT OF INTRUDER PARAMETERS Principle S Determination of relative altitude Upon confirmed transponder reception, the TCAS starts to interrogate the intruder. Its altitude is transmitted directly in the reply (standard barometric altitude) and this information is used to determine the relative altitude of the two aircraft, by calculating the barometric altitude difference. This computation is, however, only possible with respect to Mode C or Mode S transponder equipped aircraft. S Range measurement The range is calculated by measuring the elapsed time between transmission of the interrogation signal and return of the reply transmitted by the intruder. Aircraft are detected from a minimum range of 14 NM. S Determination of Azimuth There are several methods for calculating the angle of reception of a radio− electric signal with respect to a reference direction such as the aircraft centerline. The technique used in the TCAS II computer is the interferometer system. The interferometry principle is based on a comparison of signal phases received by four independent elements of the directional antenna, taken in pairs. With two poles, E2 and E4, the phase difference of signals received on the two elements depends on the angle of reception of these signals since the difference in distance to the source of the two reception poles varies with this angle. B : source signal reception angle S2−S4 S1−S3 d : phase difference of signal received on poles E2 and E4 : phase difference of signal received on poles E1 and E3 : distance between the two poles B S2 − S4 = Arc tan −−−−−−−−− S1 − S3 HAM US/F-4 SaR Dec 2005 S Tracking Once identified, the intruders are tracked by a series of interrogation−replies in Mode C only all−call for Mode C transponder−equipped aircraft, and in Mode S for Mode S transponder−equipped aircraft. These exchanges permit the TCAS to periodically update the altitude, range and bearing data for each intruder and to compute the range rate and altitude rate variations. These data are then used to determine the time separating the two aircraft from their closest point of approach. S Broadcast messages Every ten seconds, the TCAS transmits a broadcast message intended to inform nearby aircraft, themselves equipped with a TCAS, of the presence of a TCAS−equipped aircraft in their traffic area. These messages, received by the Mode S transponders, are communicated to the TCAS computer to enable it to know the number of TCAS−equipped aircraft in its detection envelope. This information is then used in the interference limitation formulas whose results modulate the Mode S interrogation output power level in inverse proportion to the number of aircraft. This reduces the number of non−elicited replies received by ground ATC stations. The messages transmitted are of the uplink format type UF = 16 with the Mode S 24 bit address of the interrogating TCAS included in the MID field (bits 65 to 88) with the UDS field (bits 33 to 40) containing the code F50. No response is expected for this type of message. S Communications frequencies Communications between two aircraft are always crossed between transponder and TCAS. The TCAS transmits at a frequency of 1030 MHz to the transponder of the other aircraft, whose reply signals are at a frequency of 1090 MHz to the TCAS receiver. This choice allows system compatibility with ground station−transponder links as the ground stations use the same frequencies as the TCAS. Page 356 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 180 HAM US/F-4 SaR Dec 2005 TCAS Intruder Parameters Page 357 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION TCAS Part-66 TCAS-Coordination Two TCAS−equipped aircraft must coordinate their maneuvers to avoid the flight path corrections ordered by each TCAS resulting in a hazardous situation. S Coordination principle In most cases of encounters between two TCAS−equipped aircraft, mutual identification is almost but not quite simultaneous, with sufficient time lag to establish the priority necessary for the coordination process. The first aircraft to detect a potentially dangerous configuration computes a deviation maneuver sense and communicates it to the other aircraft.This aircraft takes the information into account and in turn computes a correction. If two aircraft detect each other at exactly the same time and simultaneouusly transmit coordination messages containing incompatible deviation senses it is, by convention, the aircraft that has the highest Mode S address that cancels its trajectory correction. A time delay in the display of orders on the display units avoids opposing orders. S Communications protocol Communications between two aircraft comprises three phases: − Detection phase The TCAS receives squitter messages transmitted by the transponder of the intruder aircraft. These are DF11−type messages transmitted periodically at one second intervals and intended to enable its detection and identification. This message essentially uses two fields: −an AA (Address Announced) field containing the Mode S address which must be acknowledged −A PI (parity interrogator Identity) field which is the result of a parity encoding using a method of polynominal multiplication. On reception, the Address Announced field is multiplied by the same poly−nominal and the result is compared with the received value. HAM US/F-4 SaR Dec 2005 − Surveillance phase After detection, a surveillance phase starts. This phase is composed of an acquisition part and a tracking part. 1) Acquisition interrogation When the TCAS receives a squitter and acquires the Mode S address of the intruder it enters into contact by transmitting a UF 0−type message (Short special surveillance interrogation) with the following specific fields: −bit 9, RL = 0 : reply message length requested short (56 bits) −bit 14, AQ = 1 : acquisition−type message indication. 2) Acquisition reply The intruder’s transponder replies to this request by a DF 0 message containing the following information: −bit 6, VS : Vertical Status, = 1 if the aircraft is on the ground,= 0 if the aircraft is airborne −bits 9 to 11, SL : Sensitivity level indicates in which sensitivity level its TCAS is operating −bits 14 to 17, RI : combinations of bits, from values 8 to 15, specify the maximum speed the aircraft can reach. The other combinations are not used −bits 20 to 32, AC : aircraft altitude code indicating the barometric alti tude. 3) Tracking interrogation After its acquisition, the intruder is tracked by UF 0−type interrogations with the following field values: −RL = 0 : reply message length requested short −AQ = 0 : not an acquisition message. 4) Tracking reply The intruder’s transponder replies with a DF 0 message indicating altitude and TCAS sensitivity level by a combination of fields SL and RI: −SL : bits 9 to 11 −RI : bits 14 to 17, combination values 0 to 7. Values 8 to 15 are not used −AC : bits 20 to 32, aircraft altitude code. Page 358 Part-66 FORMAT No. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 181 HAM US/F-4 SaR Dec 2005 TCAS Coordination DF11 / UF0 / DF0 Page 359 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION TCAS Part-66 − Coordination phase If the intruder becomes a threat, the TCAS programs a deviation maneuver to avoid a risk of collision. A coordination procedure is initiated between the two aircraft with an exchange of the following messages : 1)Coordination interrogation The TCAS transmits a UF16 Long Special Surveillance message whose fields contain the following indications: −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | BITS | FIELD | INDICATION | |−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− −− | | 9 | RL |= 1 : reply message length requested long | | 14 | AQ |= 0 : non−acquisition type interrogation | | 33−40 | UDS |U Definition Subfield − defines the other data in the | | | |MU field(Comm−U), composed of bits 42 to 88 | | 42 | MTB |Indicates multiple threat processing | | 43−44 | CVC |Cancel Vertical resolution advisory Complement − | | | |used to cancel an RA complement sent earlier to | | | |an intruder | | 45−46 | VRC |Vertical Resolution advisory Complement − used to | | | |transmit an RA vertical complement to the intruder | | | |requesting it not to modify its trajectory | | | |(don’t climb, don’t descend) | | 47−49 | CHC |Cancel Horizontal resolution advisory Complement− | | | |not used in TCAS II | | 50−52 | HRC |Horizontal Resolution advisory Complement − not | | | |used in TCAS II | | 53−55 | |not used | | 56−60 | HSB |Encoded Sense bits for Horizontal resolution | | | |advisory complement − not used in TCAS II | | 61−64 | VSB |Encoded Sense Bits for Vertical resolution advisory | | | |complement − parity code to protect the 4 vertical | | | |command bits (43−46) | | 65−88 | MID |TCAS−equipped aircraft Mode S address. | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− HAM US/F-4 SaR Dec 2005 2)Coordination reply After acquisition of this message, the intruder’s transponder replies with a Long Special Surveillance DF 16 type message, containing the information previously transmitted to it by its own TCAS: −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | BITS | FIELD | INDICATION | |−−−−−− |−−−−− |−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | |6 | VS |Vertical Status−indicates aircraft on ground or airborne | | 9−11 | SL |With RI, SL indicates the sensitivity level at which the | | | |interrogated aircraft’s TCAS is operating | | 14−17 | RI |Reply Information | | 20−32 | AC |Altitude Code −contains aircraft altitude | | | |encoded in 100 ft increments if bit 28 equals 0, | | | |and in 25 ft increments if bit28 equals 1 | | 33−40 | VDS |V Definition Subfield defines the contents of the data and | | | |coding in the field MV (CommV) composed of bits 41 to 88 | | 41−54 | ARA |Active Resolution Advisory | | | |− indicates the RAtype currently generated by the TCAS | | 55−58 | RAC |Resolution Advisory Complement | | | |−Indicates the RA complement type currently received | | | |from other TCAS−equipped aircraft | | 59−88 | |Not used | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− Page 360 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 UF FORMAT No For Training Purposes Only DF FORMAT No Figure 182 HAM US/F-4 SaR Dec 2005 TCAS Coordination UF16 / DF16 Page 361 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 Principles of Computation In the TCAS, target aircraft are categorized depending on specific criteria varying in function of altitude. The TCAS essentially uses two types of information to perform this classification: S the relative altitude between two aircraft, known by the difference of their standard barometric altitudes S the distance or range separating them. Acquisition of these two parameters at regular intervals (tracking) enables their variations to be calculated: S altitude rate S range rate. Assessement of the potential threat represented by an intruder depends on two criteria determined with respect to a point in the traffic area called Closest Point of Approach (CPA). This is the point of minimum distance between the two aircraft, assuming that their trajectories do not deviate. The two criteria are: S vertical separation at CPA S time left before reaching CPA. − Projected vertical separation The threat is evaluated by calculating the vertical separation between the two aircraft at the closest point of approach. − Time to intercept (TAU) The TCAS does not need to locate the CPA in space, but rather it needs to know the time to intercept for two aircraft. For example, if two aircraft are approaching on the same axis on a collision course, this time is the ratio of distance between them to the sum of their speeds. More generally, the TCAS uses range and range rate measurement to compute this time : RANGE TAU = −−−−−−−−−−−− RANGE RATE R(NM) i.e. TAU(s) = 3600 −−−−−−−− RR (Kts) With the risk of collision being in inverse proportion to this time, trajectory correction orders are initiated by crossing predetermined time thresholds whose values depend on the altitude layer in which the aircraft is located. This method of calculation avoids the initiation of corrections if, from a certain distance, the TAU trend is inverted even though the distance separating the two aircraft decreases. For example, in the case of two aircraft moving on parallel axes but in the opposite direction. For Training Purposes Only The TCAS computer processes the current altitude and altitude rate of the intruder to predict whether it will be within limits considered dangerous at the closest point of approach. Intruder 1 penetrates the altitude zone delimited by the upper and lower thresholds but when it reaches CPA, it will be outside this zone. Therefore no advisory will be issued. However, when intruder 2 reaches CPA it is still inside this zone and therefore an advisory will be issued. HAM US/F-4 SaR Dec 2005 Page 362 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 183 HAM US/F-4 SaR Dec 2005 TCAS Vertical Separation / TAU Page 363 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 DEFINITION OF TARGET AIRCRAFT Target aircraft are divided into four categories : − OTHER − PROXIMATE − TRAFFIC ADVISORY (TA) − RESOLUTION ADVISORY (RA). S OTHER aircraft Aircraft detected by the TCAS are defined as OTHER if they do not enter in PROXIMATE, TA or RA categories. They are displayed on the ND in: −above configuration from −2700 ft to +9900 ft, For Training Purposes Only −below configuration from −9900 ft to +2700 ft. The THRT/ALL/ABV/BLW switch on the ATC/TCAS control unit selects these configurations. S PROXIMATE aircraft Targets are defined as proximate traffic if the difference between their altitude and that of the TCAS aircraft is less than 1200 ft and if their range is within 6 NM. Their presentation on the ND is conditioned by the presence of another TA or RA intruder (THREAT TRAFFIC function). Generally aircraft not in the immediate vicinity enter into this category. Depending on their trajectory, they may : −conserve this status and move away without an advisory being declared. In this case the pilot is informed of their presence on the ND by a white filled diamond symbol and can monitor their progress, or −have a trajectory liable to lead to a conflict situation and in this case they require a traffic advisory and their symbol changes. S TRAFFIC ADVISORY aircraft When an intruder is relatively near but does not represent an immediate threat, the TCAS issues a traffic advisory. Its presence is displayed on the ND by an amber filled circle. Its display is accompanied by an aural alert: ”Traffic Traffic”. The pilot is therefore aware of its presence and knows its range and relative bearing. Its display is linked to vertical separation and time TAU before CPA values. Depending on its trajectory, an intruder may conserve this status and move away, or it may become a collision threat. In this case avoidance maneuvers are suggested to the pilot via a resolution advisory. S RESOLUTION ADVISORY aircraft In resolution displays, the intruder is represented on the ND by a red filled square and corrective orders are issued on the vertical speed scale of the PFD. Crossing into resolution advisory occurs for a TAU time threshold 10 to 15 seconds lower than a traffic advisory threshold. Vertical separation between the two aircraft is also taken into account for this category. There are two types of resolution advisory, in function of the vertical separation value: − Preventive Advisories − Corrective Advisories. (a)Preventive advisory In this case the vertical separation is less than a threshold S1 and greater than a threshold S2. The advisory instructs the pilot to avoid certain deviations from current vertical rate, this measure being sufficient to avoid a risk of collision. On the PFD speed scale, the forbidden values are indicated by red sectors. (b)Corrective advisory In this case, the vertical separation is lower than the threshold S2. On the vertical speed scale of the PFD, colored sectors depict avoidance maneuvers to be performed: S red sector = forbidden vertical speeds S green ”fly to” sector = a vertical speed range to be respected. Page 364 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 184 TCAS Corrective / Preventive Advisory Thresholds Page 365 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION TCAS Part-66 Aural alerts Trajectory correction or holding visual orders are accompanied by synthesized voice announcements whose level cannot be adjusted by the pilot. These announcements are generated by the TCAS computer and broadcast via the cockpit loud speakers. These messages and their meanings are described below: −”CLIMB, CLIMB”: Climb at the rate shown by the green sector on the PFD (1500 ft/min), −”CLIMB, CROSSING CLIMB, CLIMB, CROSSING CLIMB”: As above except that it further indicates that own flight path will cross through that of the intruder, −”INCREASE CLIMB, INCREASE CLIMB”: Follows a ”climb” advisory. The vertical speed of the aircraft should be increased (2500 ft/min), −”CLIMB, CLIMB NOW, CLIMB, CLIMB NOW”: Follows a ”descend” advisory when a reversal in sense is required to achieve safe vertical separation from a maneuvering intruder, −”DESCEND, DESCEND”: Descend at the rate indicated by the green sector on the PFD (−1500 ft/min), −”DESCEND, CROSSING DESCEND, DESCEND, CROSSING DESCEND”: As above except that it further indicates that own flight path will cross through that of the intruder, −”INCREASE DESCENT, INCREASE DESCENT”: Follows a ”descend” advisory. The vertical speed of the descent should be increased (−2500 ft/min), −”DESCEND, DESCEND NOW, DESCEND, DESCEND NOW”: Follows a ”climb” advisory when a reversal in sense is required to achieve safe vertical separation from a maneuvering intruder, −”ADJUST VERTICAL SPEED, ADJUST”: Reduce vertical speed to that shown by the green sector on the PFD. It can correspond to a corrective reduce climb or reduce descent. It can also represent a weakening of corrective RA. Four other aural advisories are also generated: −”MONITOR VERTICAL SPEED”: Indicates that a forbidden vertical speed range exists (red sector) and that pilot must monitor vertical speed so as not to enter this range (Preventive Advisory). −”MAINTAIN VERTICAL SPEED, MAINTAIN”: Indicates a non−crossing advisory type, maintains rate RA’s (corrective). −”MAINTAIN VERTICAL SPEED, CROSSING MAINTAIN”: Indicates an altitude crossing advisory type, maintains rate RA’s (corrective). These messages are spoken only once if softening from a previous corrective advisory, −”CLEAR OF CONFLICT”: Indicates that separation has been achieved and range has started to increase. Page 366 Part-66 THIS PAGE INTENTIONALLY LEFT BLANK For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Page 367 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 ADVISORY INHIBIT CONDITIONS In certain particular conditions, certain advisories are not generated as they could lead to the pilot adopting flight conditions that are hazardous or outside the aircraft’s performance capability. S Low altitude inhibitions Ground proximity leads to the inhibition of those advisories liable to cause a hazardous situation at this level. In decreasing altitude, these are: −below 1450 ft above ground level (AGL) inhibition of ”Increase Descend” resolution advisories (RA) −below 1200 ft AGL at take−off and 1000 ft AGL in approach, inhibition of ”Descend” resolution advisories (RA) −below 1100 ft AGL at take−off and 900 ft AGL in approach, aural alerts (TA and RA) are inhibited. S High altitude inhibition Above the altitude limit given by pin program, further climb orders are inapplicable as the aircraft performance capability does not permit them to be taken into account. ”Climb” advisories are therefore inhibited above this altitude. S Advisory inhibit discretes Three discretes are used to manage priority between : Rejection of signals from aircraft on ground Aircraft on the ground may reply to TCAS interrogations, producing an unnecessary overload in the processing and display of information. The ”Ground Logic” (aircraft declared on ground) is enabled when the own aircraft descends below 1650 ft AGL and when it climbs up to 1750 ft AGL. Intruders are declared to be on ground if they are within 360 ft from the ground when descending, and if they are within 400 ft from the ground when climbing. All on−ground intruders are displayed as non−threat traffic (white unfilled diamond). Intruders declared to be on ground can never cause proximate, traffic or resolution advisories (TA, RA). But, as the altitude transmitted by the intruder is a barometric altitude with respect to sea level, the TCAS shall process this value to convert it into height above ground level in order to compare it with the 380 ft (+/− 20 ft) threshold. NOTE: If the intruder is equipped with a Mode S transponder, the ground/ flight information is sent in the DF message. The TCAS (change 7.0 only) directly knows if this intruder is on the ground. Rejection of Non Altitude Reporting Aircraft by the own aircraft The TCAS does not display ”Non Altitude Reporting Aircraft” above 15500 ft MSL. −windshear/stall For Training Purposes Only −GPWS − G/S −and the TCAS computer. The environmental alert priorities are : windshear/stall, GPWS and then TCAS II. When TCAS II is inhibited, the TA ONLY mode is selected and the voice announcements are cancelled. HAM US/F-4 SaR Dec 2005 Page 368 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 185 HAM US/F-4 SaR Dec 2005 TCAS Airborn / Ground Definition Page 369 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION TCAS Part-66 SENSITIVITY LEVEL The notion of sensitivity level is very important in the TCAS as many of the operating modes depend on it. The TCAS separates the surrounding airspace into altitude layers. A different Sensitivity Level (SL) threshold for issuing advisories is applied to each altitude layer. The sensitivity level is decreased at low altitude to prevent unnecessary advisories in higher traffic densities such as terminal areas. Generally, the level is determined automatically by the TCAS in function of: −altitude values from the radio altimeter up to 2500 ft AGL −barometric altitude values in the 2500 ft to 48,000 ft range. TAU values corresponding to each sensitivity level indicate the TA and RA thresholds. The vertical separation thresholds at CPA also vary in function of the sensitivity level for the different types of advisory. NOTE:Level 1 corresponds to Standby Mode in which no advisory is generated. NOTE:This table indicates the threshold based on own aircraft altitude. Each aircraft altitude depends on a hysteresis: S 1.000 +/− 100 ft; S 2.350 +/− 200 ft; S 5.000 +/− 500 ft; S 10.000 +/− 500 ft; S 20.000 +/− 500 ft; S 42.000 +/− 500 ft. For example, to switch from sensitivity level 3 to sensitivity level 2, the altitude must fall below 900 ft. However, to switch from sensitivity level 2 to sensitivity level 3, the altitude value must go above 1100 ft. HAM US/F-4 SaR Dec 2005 There are two other means of modifying the sensitivity level: − selecting TA only mode on the ATC/TCAS control unit forces level 2. In this case, intruders of all types are displayed but will not be transformed into RA symbols and no vertical speed modification indications will be issued. − the ATC Mode S equipped ground stations may modify the sensitivity level of the aircraft TCAS via the uplink without, however, having the capability to force the Standby Mode. If several ground stations command sensitivity levels, the TCAS logic selects the lowest level. Definition of priority logic:First a sensitivity level based on altitude is selected. Level 2 is selected if the radio altimeter altitude is less than 1000 ft. Level 2 is also selected if own aircraft is configured such that both CLIMB and DESCEND RAs are inhibited (e.g., below 1000 ft AGL with insufficient climb performance). Level 3 is selected if the aircraft is above 1000 ft and below 2350 ft AGL. If the aircraft is above 2350 ft AGL, barometric altitude is used to select either level 4 (from 2350 to 5000), level 5 (below 10,000 ft), 6 (from 10,000 to 20,000 ft), and 7 (above 20,000 ft).ATC/TCAS control unit input is read by the TCAS computer. If the pilot has selected Automatic Mode (TA/RA), then the altitude−based sensivity level will be used in comparisons to determine the final level.From all sensitivity level commands, if any, received from ground stations, the lowest is selected.If the TA ONLY mode is selected, either manually via the control unit or by a ground station, the altitude−based sensitivity level is used for TA thresholds and the RAs are inhibited. Otherwise, the lowest of all inputs is chosen. Page 370 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 186 HAM US/F-4 SaR Dec 2005 TCAS Senitivity Level Page 371 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 INFORMATION DISPLAY The TCAS information is presented on the CAPT and F/O S NDs S PFD. Additional messages can also be presented on the EWD. For Training Purposes Only Traffic information on the ND Target aircraft are presented on the ND in: S ROSE mode or S ARC mode but not in PLAN mode. These traffic indications show the situation in the surveillance zone. The aircraft present in this zone are represented by symbols whose shape and color correspond to the type of intruders defined in the TCAS. The symbols are positioned on the ND so as to depict their relative bearing and range. Data tags are associated with intruders. These tags consist of: S two digits indicating their relative altitude in hundreds of feet S a symbol indicating whether the intruder is above (+) or below (−) the aircraft. S An arrow to the right of the symbol indicates the vertical trend (500ft/min or more) of the aircraft. Targets are symbolized according to their type: S S S S OTHER : white unfilled diamond, height 7 mm PROXIMATE TRAFFIC : white filled diamond, height 7 mm TRAFFIC ADVISORY : amber filled circle, diameter 5 mm RESOLUTION ADVISORY : red filled square, side 5 mm. The display ( in this A340 example ) only presents the eight most threatening intruders (number determined through program pins on the TCAS computer). The own aircraft is represented by the aircraft symbol at the center of the dial in ROSE mode and at the lower quarter and at the center of the image in ARC mode. A white range ring with markings at each of the twelve clock positions is placed around the own aircraft symbol at a radius of 2.5 NM. HAM US/F-4 SaR Dec 2005 Page 372 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 187 HAM US/F-4 SaR Dec 2005 TCAS Indication on ND Page 373 Part-66 The Figure 188 shows an example of the display on the ND: S These indications are only presented for the 10, 20 and 40 NM range selections. If a TA or RA type intruder is detected and the display range is at a higher scale, the following message comes into view at the center of the display, in red for RA and in amber for TA: S REDUCE RANGE If a TA or RA type intruder is detected, and the ND mode is inadequate for display, ( PLAN mode ) the following message comes into view at the center of the display in the same colors as above: S CHANGE MODE No Bearing Intruder The TCAS can detect an intruder without acquiring its bearing (for instance multipath problem or MLG down). In this case its range, relative altitude and an arrow are displayed in the TCAS area (at the bottom of the ND). The color of the display is the same as the color of the intruder symbol. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS HAM US/F-4 SaR Dec 2005 Page 374 Part-66 ROSE MODE ARC MODE For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 188 HAM US/F-4 SaR Dec 2005 TCAS ROSE- / ARC- Mode Display Page 375 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 Traffic information on the PFD Resolution advisories are represented on the vertical speed scale of the PFD by indications given in the form of a band made up of colored sectors: S a red sector represents a forbidden vertical speed range S a green sector indicates the vertical speed range the aircraft should fly in to avoid a collision threat represented by one or more intruders. Preventive advisory display Preventive resolution advisories advise the pilot to avoid vertical speeds that could lead to a hazardous situation. They are represented by one or two red sectors on the vertical speed scale. The pilot must keep the vertical speed of his aircraft outside these zones. For example, if own and intruder’s flight paths are horizontal and cross through each other, the intruder being at a higher flight level, if the vertical separation is sufficient it is not necessary to modify the aircraft flight path. The positive vertical speed sector is in red to indicate that the aircraft may remain at its present level or may descend but must not climb. If intruders are detected above and below, two red sectors are displayed leaving an uncolored zone around the zero value on the vertical speed scale advising the pilot to maintain the aircraft at its current level. For Training Purposes Only Corrective advisory display Corrective resolution advisories are displayed to advise the crew to perform an avoidance maneuver in the vertical sense. This maneuver may take different forms: S climb or descent if the aircraft is in level flight S reducing or increasing rate of climb or reversing to descent if the aircraft is in climb S reducing or increasing rate of descent or reversing to climb if the aircraft is in descent. When resolution advisories are displayed, the vertical speed scale surface changes from trapezoidal to rectangular. The grey background is replaced by green and red sectors defining the optimum vertical speed values. The pilot’s task is to maneuver the aircraft to keep the needle out of the red sectors and place it in the adjacent green ”Fly−to” sector. The vertical speed information needle and digits is colored in red when the vertical speed is in the forbidden area. It becomes green when the vertical speed is in the authorized area. HAM US/F-4 SaR Dec 2005 Page 376 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Figure 189 HAM US/F-4 SaR Dec 2005 TCAS Indication on PFD Page 377 Part-66 The Figure 190 ries. shows several examples of corrective and preventive adviso- For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS HAM US/F-4 SaR Dec 2005 Page 378 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION TCAS Part-66 DESCENT CORRECTIVE ADVISORY CLIMB CORRECTIVE ADVISORY Figure 190 HAM US/F-4 SaR Dec 2005 PREVENTIVE ADVISORY TCAS PFD Indications Page 379 Part-66 Messages announced on the ND As well as intruder information, the ND also displays operating mode messages or fault data. This information is presented in the lower section of the ND (message zone): S TA ONLY − white − for the TA mode (automatic or manual switching) S TCAS − red − to indicate a TCAS computer failure. Messages announced on the PFD A red TCAS flag appears to the left of the vertical speed scale on the PFD if the TCAS cannot deliver RA data. Messages announced on the EWD S If a TCAS failure is detected, the amber warning message NAV TCAS FAULT is displayed on the EWD. S Selection of the TCAS STBY mode on the ATC/TCAS control unit results in the display of the TCAS STBY message (green) in the memo section of the EWD. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS HAM US/F-4 SaR Dec 2005 Page 380 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION TCAS Part-66 For Training Purposes Only WHITE „TA ONLY“ MESSAGE RED TCAS FLAG Figure 191 HAM US/F-4 SaR Dec 2005 TCAS Messages on ND/PFD/EWD Page 381 Lufthansa Technical Training M13.4 NAVIGATION TCAS Part-66 SELF TEST A quick check of the correct operation of the TCAS installation can be performed by activating the TEST function : S either by pressing the pushbutton switch on the front of the TCAS computer S or through the Central Maintenance System (CMS) by applying the procedure TCAS Functional Test on the Multipurpose Control and Display Unit (MCDU). The self−test sequence checks the main functions of the computer and transmits to the displays: S resolution advisory characteristics (0 ft/mn advisory, up corrective advisory, don’t descend, don’t climb > 2000 ft/mn, rate to maintain) on label 270 S label frames 130, 131, 132 containing three intruder data according to the following table: −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | INTRUDER |TYPE | RANGE | REL ALT | BEARING| VERTICAL RATE | | | (NM) | (FEET) | (DEG) | | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−| | 1 | RA | 2.00 | −1000 | +90 | no vertical rate | | 2 | TA | 2.00 | −200 | −90 | climbing | | 3 | PROX | 3.625 | +200 | +33.75 | descending | | 4 |OTHER| 3.625 | +1000 | −33.75 | no vertical rate | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− For Training Purposes Only ND image The ND must display the images corresponding to the four types of intruders : Other, Proximate, TA and RA. The shapes and colors of the traffic symbols are: S white unfilled diamond for Other traffic S white diamond for Proximate traffic S yellow circle for TA traffic S red square for RA traffic. Failure indication At the end of the test sequence the system generates a synthesized voice message: TCAS SYSTEM TEST OK if the system operates correctly or: TCAS SYSTEM TEST FAIL if an anomaly has been detected. In this case the NDs, the PFDs and the EWD show the failure messages described in Figure 191 . In addition, two windows on the front of the TCAS computer display a code identifying the failed component according to the table below: −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− | CODE | COMPONENT | |−−−−−−− |−−−−−−−−−−−−−−−−−−−−−−−−| | TP | TCAS PROCESSOR | | AT | TOP ANTENNA | | AB | BOTTOM ANTENNA | | X1 | MODE S TRANSPONDER 1 | | X2 | MODE S TRANSPONDER 2 | | RA | RADIO ALTIMETER 1 AND 2 | | PT | PITCH ATTITUDE DATA | | RL | ROLL ATTITUDE DATA | | HD | HEADING DATA | | RD | RA DISPLAY 1 AND 2 | | PP | PROGRAM PINS | −−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−−− PFD image At the beginning of the test sequence, green and red sectors must appear sequentially on the vertical speed scale of the PFD. Then a resolution advisory display is shown. HAM US/F-4 SaR Dec 2005 Page 382 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION TCAS Figure 192 HAM US/F-4 SaR Dec 2005 TCAS - Test Display Page 383 Part-66 WEATHER RADAR SYSTEM GENERAL PRINCIPLE OF OPERATION Information on the exact location and severity of bad weather is necessary to prevent diversions or cancellations of flights. What is required is an airborne system capable of detecting the weather conditions leading to the hazards of turbulence, hail and lightning. Attention has been concentrated on developing systems which will ’detect’ turbulence. If an aircraft passes through regions of severe turbulence it is obviously subject to mechanical stress, which may cause damage, possibly leading to a crash. On commercial flights, passenger comfort is also important. Unfortunately the phenomenon of rapidly and randomly moving dry or clear air currents is not amenable to detection by todays technique.As a consequence of the inability at present to detect the turbulence directly, systems have been developed which detect either water droplets (Weather Radar) or electrical activity (Storm Scope), both of which are associated with convective turbulence in cumulonimbus clouds. Clear air turbulence (CAT) has no detectable associated phenomena which can give a clue to its presence. To detect water droplets or raindrops, a conventional primary radar is used with optimized frequency, pulsewidth and power output. The maximum diameter of a water droplets which can remain in a cloud without falling to earth is dependent on the speed of the up−draught of air. If a small volume of the cloud results in strong signals from one part and weak from an adjacent part we have a steep rainfall gradient’, most probably due to a down− and up - draught close to one another. The region around this vertical wind shear is likely to be highly turbulent. Weather radar operation depends on three facts: S precipitation scatters r.f. energy; S the speed of propagation of an r.f. wave is known as the speed of light; S r.f. energy can be focused into a highly directional beam. Pulses of r.f. energy are generated by a transmitter and fed to a directional antenna. The r.f. wave will be scattered by precipitation in its path, some of the energy returning to the aircraft as an echo. The elapsed time between transmission and reception is directly proportional to range R, in particular For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 R=cxt/2 where S c is the speed of propagation (= 162 000 nM- or 300 000 Km- per second); S t is the elapsed time; and S the divisor 2 is introduced since travel is two−way. The direction of the target is simply given by the direction in which the beam is radiated. One radar mile is the time taken for an e.m. wave to travel 1nM and back, approximately 12.36 s. The pilot needs to observe the weather in a wide sector ahead of the aircraft the antenna is made to sweep left and right repetitively. Any storm cloud within the sector of scan will effectively be sliced by the beam so that a cross−section of the cloud is viewed. Display of three quantities for each target is required: range, bearing and intensity of echo. A plan position indicator (PPI) is used for display, or in EFIS equipped airplanes, the weather information is combined with the EHSI or ND indication. Page 384 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Part-66 CLOUD FORMATION SELECTED RANGE REFLECTED ENERGY (ECHO) TRANSMITTED For Training Purposes Only ENERGY PULSE Figure 193 HAM US/F-4 SaR Dec 2005 WEATHER RADAR SYSTEM Page 385 Part-66 AIRCRAFT EQUIPMENT The aircraft equipment comprises: S WX - transceiver S plan position indicator or EFIS S Control Panel S wave guide S parabolic or flat plate antenna For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 386 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Part-66 For Training Purposes Only Indicator and Control Panel Figure 194 HAM US/F-4 SaR Dec 2005 Aircraft Equipment Page 387 Part-66 COMPONENTS LOCATIONS The weather radar transceiver is located in the pressurized area of the cabin close to the forward bulkhead, to keep the waveguide length as short as possible to reduce losses. PPI or EFIS is in the cockpit as well as the control panel. The antenna is under the nose radome, which is manufactored out of glassfiber to give passage to electromagnetic waves. In some smaller business twins the antenna is located inside of the leading edge of one wing. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 388 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Figure 195 HAM US/F-4 SaR Dec 2005 Components Locations Page 389 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 WX TRANSCEIVER (HIGH POWER) The weather radar helps the pilot to detect turbulences connected to precipitation in time to be able to circumfly them. Many weather radars allow as well ground mapping for navigational purposes. The elder WX Generations are analog systems with high power outputs. The clock or the frequency of the aircraft a/c power controls the radar transmitter by switching on the high frequency generation for about 2,5 (5) sec with a repition rate of about 400 (200) Hz or PRT of 2,5 (5) msec. So the duty cycle will be 1:1000. The transmitter frequency is 9375 + 40 MHz in the X−band and the pulse power is about 50 kW to 60 kW. Passing a dublexing device and a wave guide, the transmitting pulse enters the antenna and is radiated. The beam angle is around 3o (pencil beam). The antenna is inside the radom and is turned around or left and right about 15 times per minute. Further the antenna is stabilized in pitch and roll by servo system per signal from a vertical gyro or a platform. To get weather views into altitudes above or below the flight altitude the antenna may be tilted up to + 15o. The antenna itself is either a horn radiator and a parabolic dish or slot antennas in a plate ( Flat Plate Antenna ). The echo signals from the antenna are fed into the receiver and finally are displayed on the cathode ray tube. The dublexer prevents the high energy transmitting pulses from entering the receiver. The received pulses are displayed largely with constant intensity in spite of different distances of the single aims. This is done by automatic sensitivity time control (STC) of the receiver controlling sensitivity down following each radiated transmitter pulse followed by time increasing sensitivity program. A penetration compensation increases the gain of the IF amplifier after a target has been detected, to compensate the attenuation of the electromagnetic wave inside this cloud to display also a possible second cloud with the same appropriate intensity. Modern equipment has digital memories to store the information for display on the cathode ray tube. The black/white display is replaced to day by a color tube. Figure 196 HAM US/F-4 SaR Dec 2005 WX-System Block Diagram Page 390 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 197 HAM US/F-4 SaR Dec 2005 WX Radar Transceiver Blockdiagram Page 391 Part-66 WX TRANSCEIVER (LOW POWER) The weather radar enables detection and localization of the atmospheric disturbances in the area defined by the antenna scanning: plus or minus 90 deg. of aircraft centerline and up to 320 NM in front of the aircraft. The modern WX Generations are digital systems with low power outputs. In addition the weather radar system enables: − detection of turbulence areas caused by the presence of precipitations − presentation−of terrain mapping information by the combination of the orientation of the radar beam and of the receiver gain. Five color displays are used to show precipitations and turbulences to the crew. This system is associated to: − the Inertial Reference Units for the attitude information − the Electronic Flight Instrument System (EFIS) for the generation of the distance scales and the display of the radar images. Modern semiconductor- and digital-technology allows to reduce the output power to about 125W. Control panel information is transmitted via ARINC 429 bus, and the weather radar data transmission to the PPI or EFIS is done by ARINC453 with 1600bits datawords and 1MBaud. Frequency : Wavelength: Peak Power Emitted: Range of the System: 9345 MHz X−Band 3 cm 125 W approx. 320 NM The disturbances are shown with different colors: BLACK: less than 1 mm/h Precipitation Rate GREEN: from 1 to 4 mm/h YELLOW: from 4 to 12 mm/h RED: 12 mm and above MAGENTA represents TURBULANCE AREAS up to 40 NM caused by the presents of PRECIPITATIONS at a rate of >5 m/s. Note: There is no detection of turbulences in clear sky! Note : No pre−heating time is necessary for the operation of the weather radar transceivers. CAUTION Before selecting the WX/TURB or MAP mode on the control unit, make sure that: − no one is within a distance less than 1.50 m from the antenna in movement within an arc of plus or minus 135° on either side of the aircraft centerline. − the aircraft is not directed towards any large metallic obstacle, such as hangar, which is within 5 m in an arc of plus or minus 90° on either side of the aircraft centerline. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 392 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 198 HAM US/F-4 SaR Dec 2005 WX Radar Block Diagram Page 393 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 WX RECEIVER FUNCTIONS Sensitivity Time Control, STC For correct indication operation the signal strength should depend only on the characteristics of the target and not of the range. But of course the range also affects the received power. If a target with the same reflection characteristics is received from different distances, the echoamplitude will be different. The echointensity will be less for the target at greater distance. In order to solve this problem, the receiver gain is made to vary with range, being minimum at zero range and increasing thereafter (i.e. with time). This is called sensitivity time control or swept gain. The aim is to make the receiver output independent of range. Unfortunately the received power decreases as the square of the range for targets which fill the beam, but as the fourth power of the range otherwise. The range till where the STC function is possible is about 100 nM. For Training Purposes Only Penetration compensation If two or more clouds are in the same radial direction, the indication of the second cloud has to be corrected because of the powerloss in the first cloud. This task is done by a circuit called penetration compensation. This function can only be activ within the STC range, because only within this range the receiver gain can be varied. Automatic Gain Control, AGC During the time immediately before transmission the output from the receiver is noise only, since the PRF will have been chosen to avoid second−trace echoes, i.e. the time corresponding to maximum range expires well before the next pulse. The AGC circuit is gated so that the receiver output is connected to it only for a short time before transmission. The result of having such AGC is to keep the receiver noise output constant, which under normal conditions means the gain is constant. If, for any reason there is excessive noise generated or received, the gain will fall, so keeping the background noise displayed at a constant level. An alternative arrangement is to keep the receiver gain constant regardless of how the receiver output may change. This has the virtue of keeping the conditions for inversion in the contour circuit unchanging. Such preset gain is found in modern digital systems, whereas noise−derived a.g.c. is found in older analogue systems. HAM US/F-4 SaR Dec 2005 Page 394 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 AMPLIFIER GAIN STC gain profile without echos TIME STC RANGE For Training Purposes Only TxPULS ECHO 1 ECHO 2 Figure 199 HAM US/F-4 SaR Dec 2005 WX-STC / Penetration Compensation Page 395 Part-66 Contour The pilot will be interested in those regions where the precipitation is greatest. In order to make the situation clearer, those signals which exceed a certain predetermined level are inverted in monochromatic WX radar systems, so as to show the cells of heavy precipitation as dark holes within the bright ’paint’ caused by the cloud surrounding the cell. The width of the ’Paint’ around the cell is an indication of the rainfall gradient. The narrower the width the steeper the gradient, and hence the greater the probability of encountering turbulence. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 396 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 200 HAM US/F-4 SaR Dec 2005 WX Contour Page 397 Part-66 Digitizing In modern WX systems the received and amplified analog video signal is converted into digitized weather targets in a video processor. The video processor contains an analog to digital ( A/D ) converter that digitizes analog video signals, using four comparators with variable thresholds. These time variable thresholds provide the basis for STC compensation. The threshold signal level of the digitized weather target signals is adjusted with sensitivity timing control (STC), equalizing weather intensity levels from near and far targets. The overall thresholds are also adjusted to compensate for weakened radar returns from intervening weather targets ( penetration compensation ). The video processor determines the weather intensity levels of the targets and stores these weather intensity data into range bins. The weather targets from 0 to 320 nm range are digitized with a range resolution of 1/12,8 nm for each range bin. This results in a total number of 4096 range bins. Each range bin contains a three-bit digital word. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 398 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 LEVEL magenta 100 red 011 yellow 010 green 001 black 000 TIME bin ¢ 1/12,8 nm LEVEL magenta 100 red 011 yellow 010 green 001 For Training Purposes Only VIDEO SIGNAL PROFILE black 000 6 s RETURNS 6 s TRANSMISSION Figure 201 HAM US/F-4 SaR Dec 2005 18 s RETURNS 18 s TRANSMISSION WX Video Digitizing Page 399 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 Video Processor The video processor consists of an A/D converter, scan integrator, threshold circuitry, a gain control, a STC PROM, a penetration compensator and e.g. 3 range processors. It digitizes video signals from the receiver detector, provides filtering, and incorporates STC and penetration compensation. The video processor then range processes separate sets of weather−intensity data. Stronger video signals from closer weather targets are compared with higher threshold levels. Likewise, video signals from farther out weather targets are compared with low threshold levels. The STC PROM contains two sets of firmware−programmed STC thresholds. One set is for the 6−microsecond returns. The other set is for the 18−microsecond returns. The penetration compensator contains a weather−intensity factor scale that correlates attenuation levels with storm intensity. From digitally filtered weather−intensity data feedback from the scan integrator, it determines path loss due to attenuation from intervening targets. Through the threshold circuitry, the attenuation levels lower the threshold voltage level of the A/D converter to compensate for reduction in signal due to intervening targets. The A/D converter converts the analog video signal into a three−bit digital word, representing the digitized weather intensity data. The scan integrator samples and stores this digitized weather intensity data. Each new sample is combined with the weather target returns for the same range bin from the previous transmission. This integration improves the signal−to-noise ratio of the processed weather intensity data. The scan integrator contains processed weather intensity data from 0 to 320 nm range with a range resolution of 1/12,8 nm. Each one of the range processors is dedicated to one 453 data bus. A range processor processes weather intensity data into 512 range bins at the specific radar range, based on the range selection received on the corresponding 429 control bus. HAM US/F-4 SaR Dec 2005 S Range 40 nm If the selected range is 40 nm, a range processor uses the 512 range bins of weather intensity data from the scan integrator. With the integrator’s range resolution of 1/12,8 nm, the 512 range bins correspond to 40 nm. S Range < 40 nm For range selections less than 40 nm, the range processor repeats the data, obtained from a range bin of the scan integrator, and temporarily stores them in multiple locations. If the selected range is 10 nm, the range processor receives 128 range bins of data from the scan integrator and uses them four times to compose the 512 range bins of data for transmission on a 453 data bus. S Range > 40 nm For selected ranges above 40 nm, the range processor averages multiple range bins of data from the scan integrator. If the selected range is 320 nm, the range processor averages eight consecutive range bins of data to make up one range bin of data for transmission on a 453 data bus. Video Memory Timing and control circuits control the operation of video memories 1, 2, and 3, which are dedicated to 453 data buses 1, 2, and 3, respectively. While 512 range bins (512 x 3 bits) of range processed weather intensity data and 64 bits of control data are written into a video memory to make up the 1600 bits of a display data word, three 453 bus drivers transmit the 1600−bit display data words on three 453 data buses to three indicators or display systems. Page 400 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 RANGES INPUT VIDEO 453 DATA BUS PROCESSOR #1 RECEIVER OUTPUT #2 #3 For Training Purposes Only ANT POSITION Figure 202 HAM US/F-4 SaR Dec 2005 WX Video digitizing-Schematic Page 401 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 WX CONTROL FUNCTION The weather radar system control is provided by two 32−BIT control words assembled in ARINC 429 format and transmitted over the control bus from the indicator or control unit. The R/T unit derives its operating mode, specific antenna tilt angle and other control information from a serial 32−bit control word 1, and selected range information from 32−bit control word 2. The R/T unit accepts a serial 32−bit control word only if it has the proper identifying octal Label. These are octal number 270 for control word 1 and octal number 271 for control word 2, as specified by ARINC 708. Return−to−zero (RZ) bipolar modulation is employed in conveying the one (1) and zero (0) bit logic levels as specified in ARINC 429. CONTROL WORD number 2 provides RANGE switch position and is assembled as follows: − Bits 1 through 8 makes up the label, − Bits 9 and 10 are reserved, and bits 11 through 23 are spares. − Bits 24 through 29 contain range switch selection. − Bits 30 and 31 sign/status provide matrix information VALID and INVALID DATA. − Bit 32 is set for odd parity CONTROL WORD CONTROL WORD number 1 contains GAIN, MODE, TILT and STAB control position as follows: − Bits 1 through 8 make up the label and identify the control word number of the control word appearing on the system data bus. − Bits 9, 10, 11 are reserved. − Bit 12 is for IDNT which causes the system to suppress ground clutter. − Bit 13 is for STAB which places the antenna in a stabilized mode of operation. − Bits 14, 15 and 16 identify mode selected (WX/T, TEST, WX, TURB, or MAP). − Bits 17 through 23 provide data for antenna TILT + 15 degrees in .25 degree increments. − Bits 24 through 29 provide data for controlling system GAIN. − Bits 30 through 31 provide sign/status matrix information (VALID DATA, TEST and INVALID DATA). − Bit 32 is set to give odd parity. HAM US/F-4 SaR Dec 2005 Page 402 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 203 HAM US/F-4 SaR Dec 2005 ARINC 429 Control Word Page 403 Part-66 DATA WORD The weather radar indicator display is generated from the information provided in a 1600−bit serial word, sent from the receiver−transmitter (r/t). The word consists of a 64−bit control word and a 1536−bit (return) data word. The indicator or a display management computer processes the 1600−bit serial word, stores it in a memory circuit, retrieves the data when it is ready to be displayed, then decodes and displays the return data. The control word consists of system status and scan angle components. The system status bits are used to address the memory locations of their corresponding messages. The scan angle bits are sent from the r/t to the indicator in a rho, theta coordinate system, and converted in the indicator to an X, Y coordinate system. The converted scan angle coordinates are used as address coordinates for the return data. The 1536 bits of return data are arranged in 512 3−bit words called bins. The level of rf signal reflected by the target determines the binary weighting of each bin and this is then converted by the indicator to the appropriate color signal. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 404 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 204 HAM US/F-4 SaR Dec 2005 ARINC 453 WX-Data Word Page 405 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Part-66 WX−RADAR OPERATION Radar operation is done with the controls on the indicator and/or control panel. Possible switchable operative functions are: S NORMAL S CONTOUR S CYC−lical contour: S MAPPING S TEST S ANT−TILT S GAIN S TRACE−ADJUST S ERASE−RATE S RANGE & MARKS S AZImuth S SECtor of SCAN S FRZ (Freeze) or HOLD S TGT (Target) ALERT S STAB ON/OFF The antenna is tilted by a servo system up to + 15. In MAP mode the gain of the received pulses is varied. The control (GAIN) is opened until the wanted aim is clearly visible. Are there no echoes gain is opened until the first grass appears. In weather modes always MAX. GAIN or AUTOMATIC GAIN CONTROL has to be implemented. Brightness control for setting the display brightness. (Also named Brilliance or Intensity control). The selectable distances and the metering circles interval may be read from the inscriptions. Azimuth lines 30 apart are added to the display. Aims in a sector of about + 10 between 60 and 150 NM trigger an alarm even if other distances are selected on the display or other information is displayed. In normal operation the gyro stabilization of the antenna is switched on to keep the antenna in its plane regardless off roll or pitch movements. Switched off will be only if the gyro does not work properly. HAM US/F-4 SaR Dec 2005 Page 406 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Figure 205 HAM US/F-4 SaR Dec 2005 PPI and Control Page 407 Part-66 WX ANTENNA STABILIZATION Tilt and Stabilization A weather radar may scan up to 300 nautical miles ahead of the aircraft within azimuth scan angles of typically 90. If the beam is not controlled to move only in or above the horizontal plane, part or all of the weather picture may be masked by ground returns. Imagine the aircraft rolling with left wing down. If the swept region is in the same plane as the aircraft’s lateral and longitudinal axes, and the antenna is to left, the beam will be pointing down towards the ground, while when to right, the beam will be pointing up, possibly above the weather. In fact stabilization holds the beam not in the horizontal plane but at a constant elevation with respect to the horizontal. This constant elevation is determined by the tilt control as set by the pilot. Two types of stabilization are in use: Single- and Split- Axis-Stabilization. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 408 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 azimuth fwd right rear left fwd For Training Purposes Only azimuth azimuth azimuth Figure 206 HAM US/F-4 SaR Dec 2005 PITCH ROLL SUMM SIGNAL PITCH + ROLL Antenna Stabilization Page 409 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 207 HAM US/F-4 SaR Dec 2005 Single Axis Stabilization Page 410 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 208 HAM US/F-4 SaR Dec 2005 Split Axes Stabilization Page 411 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 INDICATIONS In older and simpler radars the CRT screen is coated with a long−persistence phosphor which continues to glow some time after the electron beam has passed on its way thus storing the information over the scan interval. The modern approach to information storage is to digitize the signal which is then stored in an intermediate memory at a location depending on the scanner azimuth angle and the time of arrival measured with respect to the time of transmission.The PPI time−base scan format can be either rho−theta or X−Y as per a conventional television. In either case memory can be read at a much higherrate than that at which information is received. With many more pictures per second painted we have a bright flicker-free display. Use of the television−type display makes multiple use of the weather radar indicator relatively simple so we find such indicators able to display data from other sensors (e.g. Area nav.) or alphanumeric data such as checklists. The latest development combines the weather radar indication with EFIS and the targets are adequately shown in the EHSI or ND.Digitizing the signal involves the recognition of only discrete values of signal intensity. The standard practice is to have three levels of non−zero intensity,the highest corresponding to the contour inversion level. Although these three levels correspond to different degrees of brightness of the paint, with the use of a colour tube they can be made to correspond to three colours. Red is the choice for contourable targets. With a suitable choice of minimum signal level for a ’paint’ uncorrelated noise is virtually eliminated from the display. HOLD: Nomentarily pressing the Hold pushbutton freezes the image displayed on the indicator. Target updating will return when the hold pushbutton is pressed a second time. Selection of hold mode causes the HOLD message to appear in the center top portion of the display, alerting the pilots of the mode selected. Offset Display Controls: Momentarily selecting LFT causes a section of the antenna 180 scan to display the area ahead and to the left of the airplane with the origin (aircraft nose) at the bottom right of the display. If RGT is selected, the indicator displays the scanned area directly ahead and to the right of the airplane with the origin at the bottom left of the display. Momentary selection of FWD view selects the scanned area directly ahead and 66 percent of the range to either side with the origin at the bottom center of the display. The indicator automatically powers on with the forward view displayed. HAM US/F-4 SaR Dec 2005 WX: Pressing the weather (WX) pushbutton enables the receiver − transmitter and will cause the weather map to be displayed. WX/T: Pushing this button activates a display of detected precipitation and turbulence within 50 N miles. TURB Pressing TURB displays detected turbulence within 50 N miles. MAP: Pressing the MAP pushbutton enables the receiver − transmitter and causes the ground map to be displayed. TEST: The test mode is selected by depressing the TEST pushbutton. After a 1−second delay, a set of concentric arcs will be displayed on the indicator. The bottom arc is green, followed by yellow, red and magenta. SPRS: The suppression pushbutton activates the clutter suppression mode to reduce ground clutter in the WX mode. Any mode or range change deactivates the clutter suppression mode. RANGE: Momentarily pushing any of the range pushbuttons selects that range as the maximum displayed range nautical miles. STAB: Stabilization is engaged by depressing the STAB pushbutton. Stabilized operation places the antenna under control of the receiver−transmitter to correct for changes in airplane attitude. PWR: This pushbutton controls power to the indicator and to the receiver/transmitter. Page 412 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 RANGE MARKS INTENSITY TILT INTENSITY For Training Purposes Only GAIN Figure 209 HAM US/F-4 SaR Dec 2005 Plan Position Indicator Page 413 Part-66 WX WX is the normal mode for weather detection. WX is annunciated and gain adjust is disabled. The receiver portion of the R/T is set to automatic gain control. Weather targets closer than 40 nmi may disappear because of the characteristics of the fiat plate antenna. This is because the fiat plate antenna does not receive returns from side Lobe as is the case of a parabolic type antenna. To compensate for this characteristic, Large down tilt settings are necessary in the 40 nmi range and at high altitudes. Selected range and tilt information is annunciated and SPRS, HOLD and USTB may be selected in weather modes. If any mode is selected before time delay runs out, STBY is annunciated. During standby, the antenna is eLectricaLLY Locked to 000 degrees in pitch and azimuth and the transmitter section of the R/T is disabled− WX-T This mode is used for detection of weather and with an addition of turbulence detection. Turbulence is noted by the color magenta. Turbulence is detected by measuring precipitation velocity. NOTE: Clear air turbulence cannot be detected in this mode. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 210 HAM US/F-4 SaR Dec 2005 Weather-/Turbulence- Mode Page 414 Part-66 MAP MAP mode is used in conjunction with the tilt control. The antenna must be tilted, a shorter range should be selected, and the gain control is now active. The colors generated during map mode are green and yellow, with yellow being the most intense. Other displays (EFIS) may use different colours, to distinct between WX and MAP. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 211 HAM US/F-4 SaR Dec 2005 Map Mode Page 415 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 SAFETY PRECAUTIONS There are two hazards when operating weather radar, namely damage to human tissue and ignition of combustible material. The greater the average power density the greater the health hazard. Among the most vulnerable parts of the body are the eyes and testes. The greater the peak power the greater the fire hazard. Any conducting material close to the antenna may act as a receiving aerial and have r.f. currents induced. There is obviously a risk, particularly when aircraft are being refuelled or defuelled. An additional hazard, which does not affect safety but will affect the serviceability of the radar, is the possibility of very strong returns if the radar is operated close to reflecting objects. The result of these returns is to burn out the receiver crystals which are of the point contact type. The safe distances for radars vary widely, depending on average power transmitted and beam width. To ensure safety precautions are observed consult manufacturers’ data for safe distances then, if operating the radar for maintenance purposes, place radiation hazard warning notices the appropriate distance from the nose. When working by the antenna with the radar on standby a notice should be placed by the controls stating ’do not touch’, or better still the transmitter should be disabled or the waveguide run broken and a dummy load fitted. If the radar is switched on prior to taxiing, the transmitter should not be switched on until clear of the apron. X−ray emission is a possible hazard when operating the transmitter with the case removed, such as might be done in a workshop. The likelihood of danger is small but the manufacturers’ data should be consulted. The following rules should be observed when operating the weather radar on the ground: 1. ensure that no personnel are closer to a transmitting radar scanner than the maximum permissible exposure level m.p.e.l. boundary, as laid down by the system manufacturer; 2. never transmit from a stationary antenna; For Training Purposes Only 3. do not operate the radar when the aircraft is being refuelled or defuelled, or when another aircraft within the sector scanned is being refuelled or defuelled; 4. do not transmit when containers of inflammable or explosive material are close to the aircraft within the sector scanned; 5. do not operate with an open waveguide unless r.f. power is off; never look down an open waveguide; fit a dummy load if part of the waveguide run is disconnected; 6. do not operate close to large reflecting objects or in a hangar unless r.f. energy absorbing material is placed over the radome (RAM cap). HAM US/F-4 SaR Dec 2005 Page 416 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 non−scanning Antenna scanning Antenna unknown System and For Training Purposes Only scanning Antenna Figure 212 HAM US/F-4 SaR Dec 2005 Safety Distance Page 417 Part-66 TURBULENCE / WINDSHEAR GENERAL WXR-Transceiver of the old generation using magnetron-transmitters, were only able to detect precipitation and ground returns by reflection and to indicat those on a screen. The modern generation of solid-state airborne weather radar- as a result of the digital- and processor- technology - is capable of detecting and processing turbulences and windshear conditions, which in fact are much more important for a safe flight than precipitation. A 9.33−GHz signal is transmitted to provide a return signal from precipitation. The reflected signal is processed to determine range and azimuth of the precipitation. S The intensity of the reflected signal allows conclusions to the amount of precipitation. S Additionally for selected ranges under 80−nmi the signal is processed to determine the velocity distribution of precipitation targets to determine turbulences. S For landing and takeoff situations, the return signal is automatically processed to determine the horizontal components of precipitation being pushed towards or away from the aircraft by windshear events. S The transmitted signal is also used to produce a return from terrain features that, when processed, will produce a ground map. To fulfill all this additional functions, the WXR-Transceivers must generate a very stable frequency ( crystal controlled ) and the A. PRF ( 180 till 3.000 pps ), B. Pulse Widths ( 1 till 20 µs ) and C. Bandwidth ( optimized to the pulse width ) must be able to be adopted to the different requirements. Operation Introduction The modern WXR- generation is a solid-state airborne weather radar, operating in the X- Band and indicating the weather in color displays. The system scans the area in front of the aircraft, to detect precipitations, turbulences and windshear conditions which may be hazardous for aircraft and / or passenger. Radar is an acronym for radio detection and ranging and uses the echo-principle. For airborne weather radar the reflectors are S Rain S Hail S Snow or S Ice cystals. Modern WXR - Systems generate very short pulses ( 1 bis 20 µs ) with pulse power of about 100W. The receiver looks for the echoes processes the signal and brings it into a format ( ARINC 453 ) that can be used by the EFIS or PPI to display the information on a screen. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 418 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 213 HAM US/F-4 SaR Dec 2005 Wetter Detection Page 419 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 DETECTION OF PRECIPITATION TARGETS The amount of energy reflected depends on the reflective quality of the target. When a short pulses strike a target such as precipitation, some of the energy is absorbed, some of it is refracted, and the remainder is reflected. lt has been found that heavy rainfall produces the strongest reflections for weather radar. Light rainfall, snow, and ice crystals produce weak returns. The strength of the return signal is related to the reflective properties of the target and the distance the pulse must travel. Radar systerns compensate for the attenuation of the signal due to the distance traveled, by a circuit called sensitivity time control (STC). The STC circuit controls the receiver sensitivity with respect to time and thus range. The receiver sensitivity increases during the period between transmitted pulses when the receiver is ”listening” for returned pulses. Modern digital systems additionally change the threshold levels for digitizing of the analog video, so that the weather image for the total range will be calibrated. Refer to Figure 214 for a representation of the returned signals processed by the receiver. The received signal provides strength of signal information and the range to target. The range of the target is determined by the amount of elapsed time that occurs between the transmission of a pulse and the reception of the reflected or return signal. The direction or azimuth bearing of the target is determined by noting the azimuth pointing position of the antenna. Bearing and range information is then coupled with the reflectivity information and applied to the indicator. Due to the closing speed between aircraft and target, there can be a frequency shift ( doppler shift ) between transmit - and receive - frequency. HAM US/F-4 SaR Dec 2005 Page 420 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 214 HAM US/F-4 SaR Dec 2005 Weather Return Signal Page 421 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 PULSE PROPAGATION Modern WXR Receiver−Transmitter are capable of multiple pulse widths as well as multiple PRFs. That is, the pulse width transmitted is selected based upon the selected range and mode of operation. By using multiple pulse widths, the receiver−transmitter can optimize the operation of the system for the particular selected range and operation mode. This optimization provides a better resolution in weather target display and efficient use of the transmitted energy. Another factor selected by the rt to optimize performance is the pulse repetition frequency (prf). The prf is the number of times the pulse is transmitted in a 1 second period. At shorter selected ranges, the receiver portion of the receiver−transmitter is not required to ”listen” for relatively long periods of time. So, the transmitter section is free to increase the number of transmitted pulses. The greater number of transmitted pulses will provide more information and a faster update of the weather targets, but there must be a method to clearly assign the echoes to the transmit pulses. For Training Purposes Only Short Ranges ( < 60 nmi ) At short ranges (selected ranges less than or equal to 60 nmi) four pulses are transmitted once per epoch (approximately 5.6 ms). An epoch is the time interval in which a radial of radar data is processed. The prf for this short ranges is 1280 Hz. Additionally for all selected ranges, the prf is dithered by $ 75 µs to decorrelate alien radar returns. Windshear Operation For windshear detection a high number of returns is necessary. For windshear operation the dithering between 64−pulse sets is ± 192microseconds. That is, the dithering or delay occurs in 10 microsecond increments from 0 microseconds to 384 microseconds. Dither Method The prf dithering is used to reject returns from other weather radar systems (alien) that may be operating in the area of the aircraft and/or to clearly assign the echoes to the transmit pulse. The prf rate is dithered between pulse sets at the shorter ranges and between pulse sets and pulses at the longer selected ranges. The dithering is randomly delaying the transmission of the next pulse set (or pulses) by increments of 10 microseconds from 0 (no delay) to 150 microseconds in weather and turbulence detection modes. There is no dithering between consecutive transmit pulses which are used to generate turbulence or weather data. The dither value is caIculated by using the 1 and Q parameters, the previous dither value and two scaling factors. This caIculation produces a random dither that allows the receiver to suppress alien radar returns and to assign the returns to the transmissions. Medium Ranges ( > 60 nmi bis < 165 nmi ) For medium ranges (selected ranges greater than 60 nmi and less than or equal to 165 nmi) two pulse are transmitted within each epoch of about 5,6 ms. The prf for this ranges is 360 Hz. Long Ranges ( > 165 nmi bis 320 nmi ) For selected ranges greater than 165−nmi two pulses are transmitted once per epoch (8.4 ms) at a prf rate of 240 Hz. HAM US/F-4 SaR Dec 2005 Page 422 Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Part-66 MEDIUM/ LONG For Training Purposes Only ( 5.6 ms Medium Range ) Figure 215 HAM US/F-4 SaR Dec 2005 Transmit Pulse Propagation Page 423 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 TURBULENCE DETECTION The weather radar uses the greater number of pulses to produce information about target turbulence and windshear events. Refer to Figure 216 for the frequency spectrum of the return signal. The frequency of the return signal will be offset from the transmitted frequency because of the Doppler shift caused by the velocity of the aircraft with respect to the target. In addition to the frequency shift caused by the aircraft velocity, a frequency shift caused by the movement of the precipitation. To measure the spectrum width of the frequency shifts caused by precipitation movement, the Doppler shift due to the aircraft must be filtered out. To provide an accurate spectrum of return signal frequencies, a large number of samples (returns) must be used to produce accurate and reliable results. For this reason, the prf, in the turbulence mode of operation, is increased from 240 Hz to 1280 Hz. With this large number of transmit pulses, the weather radar will process every pulse for precipitation information and provide turbulence data from the spectrum of Doppler shifts caused by precipitation movement. Because of the high prf rate, the maximum range for turbulence detection is approximately 56 miles. Once an accurate spectrum of return signals is obtained, the turbulence processing circuits must determine if the spectrum represents the spectrum of a turbulent target. A turbulent target is one that exhibits a wide variance in particle (raindrop) velocity. The broader the spectrum, the greater the turbulence. The threshold of turbulent targets is precipitation velocities of 5 m/s. The five meters/second threshold corresponds to the threshold between light and moderate turbulence that could cause food and beverage spillage or possibly minor injury. This threshold translates into a Doppler frequency shift of 312.5 Hz. Target Spectrums Frequency Spectrum Shift Figure 216 HAM US/F-4 SaR Dec 2005 Frequency spectrum Page 424 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 217 HAM US/F-4 SaR Dec 2005 Return Signal with Target Spectrum Page 425 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Part-66 WINDSHEAR DETECTION Another hazardous weather−related target is a windshear event, such as a microburst. Windshear is a condition in which the wind abruptly changes its speed or direction (or both) over a small distance. lt can be associated with frontal systems, occurring over a large area, or with thunderstorms, occurring over a small area. Windshear can occur at altitude without the presence of clouds, where it is referred to as clear air turbulence, or near ground level where it has an impact on the takeoffs and landings of aircraft. lt is the windshear occurring near the ground level, referred to as a downburst, microburst or macroburst that presents the greatest danger to aircraft. A downburst refers to one of two types of windshear: macroburst and microburst. Macroburst A macroburst is a large downburst with the horizontal component of the outburst winds extending in excess of 4 km (2.5 miles). Damaging winds, lasting 5 to 30 minutes could have velocities up to 130 mph. For Training Purposes Only Microburst A microburst is a small downburst with the horizontal component of the outburst winds extending only 4 km or less. An intense microburst could induce winds as high as 170 mph, normally lasting less than 10 minutes. HAM US/F-4 SaR Dec 2005 Both macrobursts and microbursts can cause extensive damage at ground level. The macroburst is characterized by a succession of downdrafts soft landing beneath the parent raincloud. Since the cold air dome is heavier than the warm air surrounding it, the atmospheric pressure inside the dome is higher than its environment. This causes the cold air to push outward, inducing gusty winds behind the leading edge of the cold air outflow. A downburst can be detected on a Doppler weather radar because of its characteristic velocity signature where areas of positive and negative velocity over a small distance would indicate the impact signature of the downburst. Figure 218 shows what happens to an aircraft landing with a microburst present. As the aircraft approaches the microburst, it encounters the strong headwinds from the outflow. The headwind increases the aircraft’s lift and airspeed. The increased airspeed from the headwinds, would normally cause the pilot to decrease throttle in order to maintain the descent. After the strong headwinds the aircraft enters, momentarily, the center of the microburst with a strong downflow. The aircraft then encounters the outflow again, but this time a strong tailwind decreases the airspeed and the lift of the aircraft. The aircraft has already reduced speed to maintain its flight along the glidepath against the headwind. Now with the tailwind, the aircraft loses much of its remaining lift. The pilot must respond with full throttle and increase the angle of attack to recover the lost airspeed. The danger of the microbursts are the quick transition of airspeed. lf the outflow of the microburst induces winds that are 40 mph, the aircraft must handle a sudden 80 mph change in airspeed (from a 40 mph headwind to a 40 mph tailwind). For this reason, the pilot needs as much warning as possible that a windshear event exists in the flight path. Page 426 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Figure 218 HAM US/F-4 SaR Dec 2005 Approaching a Microburst during Landing Page 427 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION WEATHER RADAR Part-66 The existence of a downburst is indicated by the condition, that within a short distance, a great variance of positive to negative Doppler frequency shift is measured. For windshear detection a higher number of transmit pulses are required to gather the windshear data, In the windshear mode of operation, the pulse pattern of the receiver−transmitter is a pulse set consisting of 64−pulses transmitted once per epoch. The duration of an epoch is about 23 ms and is the time, in which datas are collected and processed to get information about windshearconditions. The prf rate for the 64 pulses is 3000 Hz. The prf rate is dithered between pulse sets to decorrelate alien radar returns. The prf dither will vary the dither period between $ 192 µs in 10 µs increments. The scan rate of the antenna is also changed, because during windshear operation the antenna scans only a $ 60 -area (120 $) instead of the normal weather mode area of $ 90 $. During windshear operation, the S weather processing is operating during right−to−left ( CCW ) scans and S windshear processing is operating during left−to−right ( CW ) scans. Additionally, the targets detected as being windshear targets are displayed only in the area $ 30 -from the aircraft center line. The return pulse information is then used to produce a distribution of the positive and negative velocity targets that are detected. The spread of target velocities will indicate if a downburst is present. The grade of the slope of Figure 219 represents the hazard to the aircraft. A steep slope (fast transition) indicates more hazard. The hazard may also be expressed as the hazard factor F, which describes the severity of a windshear event and was developed by NASA ( not described in this manual ). HAM US/F-4 SaR Dec 2005 Figure 219 Determining Windshear Hazard Page 428 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 220 HAM US/F-4 SaR Dec 2005 Detecting Microburst with Doppler Shift Page 429 Part-66 WINDSHEAR ACTIVATION WINDSHEAR ALERT REGIONS AND LEVEL Because of the hazard a windshear event presents to an aircraft during takeoff and landings, the detection of windshear events is automatically enabled when the aircraft is taking off or landing. S Anytime the weather radar is ”ON”, the windshear detection mode is automatically activated < 2.300 ft above ground level (AGL), . S The radar is automatically turned on, when the radio altimeter reports an altitude < 2.300 ft ( FAA requires < 1.200 ft AGL ) and the aircraft is in the air or on the runway ( not in the hanger). − the determination, that the aircraft is on the runway and not at the gate or in the hangar, is done by socalled Qualifiers, which are dependent from aircraft configuration. S One source could be the transponder on signal, to determine that the gateposition was left. S That the aircraft is not in the hangar, can be determined by the engine low oil pressure switch or engine start switch. S lf the radar is already operating in a weather detection mode when a windshear hazard event is detected, no pilot intervention will be required. S lf the radar is On, but in STBY, MAP, or TEST mode, when a windshear hazard event is detected, the radar operation will automatically change to the S WX+T mode ( if selected range is < 60 nmi ) or S WX mode ( if selected range is > 60 nmi ) to display weather and windshear icons. The selected range does not change. Figure 221 shows, that windshear can be detected and indicated until 5 nmi ahead and within $30$ of aircraft heading. The detected windshear events are categorized by S location of the event relative to the longitudinal axis of the aircraft and S distance from the aircraft. Within the alert region there are three levels of alerts for windshear conditions: S ADVISORY ( is deactivated at Boeing ) The least severe level is the windshear advisory alert (level 1). The windshear advisory alert is generated whenever a detected windshear event occurs within the scan area but outside the regions where a windshear hazard event would generate a windshear warning or windshear caution alert. The output generated for a windshear advisory alert is the display of the windshear icon on the display. S CAUTION This second level of alert is the windshear caution alert (level 2). The windshear caution alert is generated whenever a detected windshear event occurs outside the windshear warning alert region but within $30$ of the aircraft heading and 3 nmi from the aircraft. S WARNING The third level (most severe) of alert is the windshear warning alert (level 3). The windshear warning alert is generated whenever a detected windshear event occurs within $ 0.25 nmi and within $30$of the aircraft heading. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR During takeoff roll, the windshear warning alert occurs within 3 nmi and for approach within 1.5 nmi. CAUTION - und WARNING - INHIBIT − Takeoff : S TAS > 100 knots until RA > 50 ft. − Approach : S RA < 50 ft − RA > 1.200 ft HAM US/F-4 SaR Dec 2005 Page 430 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION WEATHER RADAR Figure 221 HAM US/F-4 SaR Dec 2005 Windshear Alert Regions Page 431 Part-66 WINDSHEAR INDICATION The windshear warning alert is announced to the crew by: S an aural warning message, generated by the radar synthesized voice: − GO AROUND WINDSHEAR AHEAD in approach or, − WINDSHEAR AHEAD, WINDSHEAR AHEAD at takeoff, S a visual warning: − on windshear indicator lamps − red W/S AHEAD message on the PFD. The location of the windshear phenomenon is shown by an icon superimposed on the radar image of a dedicated radar display ( PPI ) or the EFIS - ND/ EHSI. This icon consists of alternating red and black arcs. Yellow radial lines appear at the edges and start beyond the windshear event. These lines, superimposed on the radar image, continue to the edge of the display area to provide directional information for the event. The windshear data are always displayed even if the selector switch on the radar control unit is set to OFF. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 432 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION WEATHER RADAR Part-66 WINDSHEAR AHEAD WINDSHEAR AHEAD or GO-AROUND WINDSHEAR AHEAD WINDSHEAR For Training Purposes Only WARNING Figure 222 HAM US/F-4 SaR Dec 2005 Windshear Indications Page 433 Part-66 WXR / WINDSHEAR - TIME SHARING When the Radar operates in Weather - and Windshear ( FLW/PWS ) - Mode, and if different Ranges are selected, the operation between between the different modes and ranges are in a time sharing sequence as shown in Figure 223 . When FLW is activated, the sweep is reduced to $ 60 $ to provide a 3 sec scan per sweep. The counterclockwise (CCW) sweeps are used to obtain the windshear data. There is no update to the display during the CCW sweeps. The update weather and windshear icons are displayed during the clockwise (CW) sweeps. Refer to Figure 223 for the time sharing sequence between indicators when FLW is activated. For Training Purposes Only Lufthansa Technical Training 13.4 NAVIGATION WEATHER RADAR HAM US/F-4 SaR Dec 2005 Page 434 Part-66 For Training Purposes Only Lufthansa Technical Training 13.4 NAVIGATION WEATHER RADAR Figure 223 HAM US/F-4 SaR Dec 2005 WXR - / FLW - Operation Time Sharing Page 435 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 GLOBAL POSITIONING SYSTEM ABBREVIATIONSLIST A−S AOD bps C/A DGPS DoD DOP FANS FOM GPS GPSSU GDOP GLONASS GNSS HDOP IRU LAAS L1 L2 L−band LOP LRU MCS NAVSTAR P−channel PCM P−code RAIM Anti−spoofing Age of Data Bits per second Coarse/Aquisition GPS signal Differential GPS Department of Defense Dilution of Precision Future Air Navigation Systems (Activity of ICAO) Figure of merit (sometimes called FM) Global Positioning System GPS Sensor Unit Geometric dilution of precision Global Navigation Satellite System Global Navigation Satellite System (GPS, GLONASS, others) Horizontal dilution of precision Inertial reference units Local Area Augmentation System GPS L−band signal 1 (1575.42 Mfiz) GPS L−band signal 2 (1227.6 Mflz) L−band frequency (about 1−2 GHz) Line of position Line replaceable unit Master Control Station (for GPS − at Colorado Springs) Navigation System with Timing and Ranging Precision code channel Pulse code modulation Precision code − provided for military GPS users Receiver autonomous integrity monitoring HAM US/F-4 SaR Dec 2005 PDOP PPS PR PRN RSS RTCA S−band SA SPS SV TDOP TT&C UERE UTC VDOP WAAS Y−code Position dilution of precision (x,y,z) Precise Positioning Service Pseudorange Pseudo random noise Root sum square Radio Technical Commission for Aeronautics Microwave frequency band, about 2−4 GHz Selective availability Standard Positioning Service Space Vehicle Time dilution of precision (t) Tracking, telemetry and control User equivalent range error Universal Time Coordinated Vertical dilution of precision (z) Wide Area Augmentation System Encrypted P−code Page 436 Part-66 THIS PAGE INTENTIONALLY LEFT BLANK For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS HAM US/F-4 SaR Dec 2005 Page 437 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 GPS GENERAL Modern aviation with its tremendous increasing flight frequencies needs a very accurate knowledge of position at any time and position. Radionavigation systems like VOR/DME, Omega or the INS don‘t fulfill the requirements in accuracy and availability of modern navaids. Already after the successful start of the first satellite ( sputnik ) in 1957, the development of a satellitenavigation-system starts. It starts with the American NNSS ( Navy Navigation Satellite System ), better known as TRANSIT. TRANSIT uses 6 satellites and calculates the position by measuring the dopplershift. The modern satnav systems ( NAVSTAR : Navigation System with Timing and Ranging ) are the american GPS (Global Positioning System) and the russian GLONASS ( Global Navigation Satellite System ). The GPS is a satellite-based radio navigation system providing world wide full time precise 3D position determining for an unlimited number of (passive service) users in all weather conditions. It was developed by NATO nations, is resistant against interference and jamming and allows common grid reference. Parallel to GPS there is another satellite navigation system existing, developed by the former Soviet Union and now under control of the CIS (commonwealth of independent states) Space Command and named Global Navigation Satellite System GLONASS. Both systems are summed under the expression Global Navigation Satellite System (GNSS). Some navigation receivers use only the GPS, others are able to process both. Because both systems originally were developed for military use, only military users receive and can use the Precise Position Service ( PPS ). The civil users receive only the Standard Position Service ( SPS ). The following treats only GPS. HAM US/F-4 SaR Dec 2005 Page 438 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 224 HAM US/F-4 SaR Dec 2005 Comparison of Navigation Accuracy Page 439 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 SATELLITE RANGING PS is based on satellite ranging. That means that we figure our position on earth by measuring our distance from a group of satellites in space. The satellites act as precise reference points for us. You might ask: ”How do we measure exactly how far we are from a satellite way out in space? And how do we know exactly where a moving satellite is?” These are a couple of the details we’ll ignore for the moment. Trust me, they can be figured out. Let’s just assume for right now that we can figure out exactly where a satellite is in space and exactly how far we are from it. Then the basic concept behind GPS is simple: Let’s say we’re lost and we’re trying to locate ourselves. If we know that we are a certain distance from satellite S1, say 18.000 km, that really narrows down where in the whole universe we can be. It tells us we must be somewhere on an imaginary sphere that is centered on the satellite and that has a radius of 18.000 km. Now if at the same time we also know that we’re 19.000 km from another satellite, satellite S2, that narrows down where we can be even more. Because the only place in the universe where we can be 18.000 km from satellite S1 and 19.000 km from satellite S2 is on the circle where those two spheres intersect. Then if we make a measurement from a third satellite we can really pinpoint ourselves. Because if we know that at the same time we’re 20.000 km from satellite S3, there are only two points in space where that can be true. Those two points are, where the 20.000 km sphere cuts through the circle that’s the intersection of the 18.000 km sphere and the 19.000 km sphere. That’s right. By ranging from three satellites we can narrow down where we are to just two points in space. (A little later we’ll see that there’s a technical reason why we have to make another measurement — but for now, theoretically, three measurements are enough). How do we decide, which one of those two points is our true location? Well, we could make a fourth measurement from another satellite. Or we can make an assumption. Usually, one of the two points is a ridiculous answer. The incorrect point may not be close to the earth. Or it may have an impossibly high velocity. The computers in GPS receivers have various techniques for distinguishing the correct point from the incorrect one. Incidentally, if you’re sure of your altitude, like mariners are (they know they’re at sea level), you can eliminate one of the satellite measurements. One of the HAM US/F-4 SaR Dec 2005 spheres in our drawings can be replaced by a sphere that’s centered at the earth’s center and has a radius equal to your distance from the center of the earth. Anyway, if we wanted to be absolutely technical, trigonometry says we really need four satellite ranges to unambiguously locate ourselves. But in practice, we can get by with just three if we reject the ridiculous solutions. And that’s it. The basic principle behind GPS: using satellites as reference points for triangulating your position somewhere on earth. Everything else about the system is just technical details designed to help carry out this ranging process — to make it more accurate and easier to do. Now let’s look at some of those details. SUMMARY S Position is calculated from distance measurements to satellites. S Mathematically we need four measurements to determine exact position. S Three measurements are enough if we reject ridiculous answers. S Another measurement is required for technical reasons to be discussed later. Page 440 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 R1 18.000 km ÄÄ ÄÄ ÄÄ ÄÄ S1 1.One measurement gives us a sphere where we are Ä Ä Ä Ä S1 For Training Purposes Only R1 ÂÂ ÂÂ ÂÂ ÂÂ ÂÂ ÂÂ ÂÂ R2 19.000 km Ä Ä Ä S3 S2 2. Two measurements gives us a circle where we are Figure 225 HAM US/F-4 SaR Dec 2005 ÄÄ ÄÄ ÄÄ R3 20.000 km 3. Three measurements gives us only two points in space where we must be GPS Satellite Ranging Page 441 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 DISTANCE MEASURING FROM A SATELLITE Since GPS is based on knowing your distance to satellites in space, we need a method for figuring out how far we are from those satellites. Surprisingly, the basic idea behind measuring a distance to a satellite is just the old ”velocity times travel−time” equation we all learned in school. You remember those word problems: ”If a car goes 60 miles an hour for two hours, how far has it gone?” Well, it’s Velocity (60 miles/hour) times travel time (2 hours) equals distance (120 miles). The GPS system works by timing how long it takes a radio signal to reach us from a satellite and then calculating the distance from that time. Radio waves travel at the speed of light: 186,000 miles per second. So if we can figure out exactly when the GPS satellite started sending its radio message and when we received it, we’ll know how long it took to reach us. We just multiply that time in seconds by 186,000 miles per second and that’s our range to the satellite. (And remember, all we need is three ranges to three different satellites and we’ve got our position). Now of course, our clocks are going to have to be pretty good with short times because light moves awfully fast. In fact, if a GPS satellite were right overhead, it would only take about 6/100ths of a second for the radio message to get to us. So, in a way, GPS is a child of the electronic revolution. The kind of timing accuracy it demands is only possible because very precise electronic clocks are now relatively inexpensive. We’re all familiar with those $20 quartz watches that keep unbelievable time. Well, GPS relies on an advanced form of that kind of timing. In fact most receivers can measure time with nanosecond accuracy. The big trick to measuring the travel time of the radio signal is to figure out exactly when the signal left the satellite. To do that the designers of the GPS system came up with a clever idea: Synchronize the satellites and receivers so they’re generating the same code at exactly the same time. Then all we have to do is receive the codes from a satellite and then look back and see how long ago our receiver generated the same code. The time difference is how long the signal took to get down to us. An everyday analogy. To picture how this works, imagine you and a friend were standing at opposite ends of a football stadium. Now suppose there was a way to make sure, that you both started counting to ten at exactly the same moment. And you both yelled the numbers out as you counted. HAM US/F-4 SaR Dec 2005 What you would hear at your end of the football stadium would be yourself saying: ”One... two.. three...” and then, a bit later, you’d hear your friend’s voice saying ”one... two...” and so on. You might already be up to ”three” by the time you heard him saying ”one.” That’s because it takes a while for the sound of his voice to get all the way across the stadium to you. Since you both stared yelling at the same time you could measure the time between when you said ”one” and when you heard your friend say ”one”. That time would be the travel time for sound to cross the stadium. That’s basically how the GPS system works. The advantage of using a set of codes, or in the case of our analogy, a string of numbers, is that you can make the time measurement any time you want. The GPS system doesn’t use numbers however. Both the satellites and the receivers actually generate a very complicated set of digital codes. The codes are made complicated on purpose so that they can be compared easily and unambiguously, and for some other technical reasons .The codes are so complicated they almost look like a long string of random pulses. They ’re not really random though, they’re carefully chosen ”pseudo−random” sequences that actually repeat every millisecond. So they’re often referred to as the ”Pseudo−Random Noise” PRN-Code. SUMMARY S The distance to a satellite is detemined by measuring how long a radio signal takes to reach us from that satellite. S We assume that both the satellite and our receiver are generating the same pseudorandom code at exactly the same time. S We know how long it took for the satellite’s signal to get to us by comparing how late its pseudo−random code is, compared to our code. Page 442 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 226 HAM US/F-4 SaR Dec 2005 Acquisition of the GPS C/A Signal Page 443 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 PSEUDO RANGE MEASURING We know that light travels at 186,OOO miles ( or 36.000km ) a second. If the satellite and our receiver were out of sync by even 1/100th of a second, our distance measurement could be off by 1,860 miles! How do we know both our receiver and the satellite are really generating their codes at exactly the same time? Well, at least one side of the clock sync problem is easy to explain: the satellites have atomic clocks on board. They’re unbelievably precise and unbelievably expensive. They cost about one hundred thousand dollars apiece and each satellite has four, just to be sure one is always working. That’s fine for the satellites, but what about us mortals down here on earth. If we had to have a hundred thousand dollar atomic clock in every GPS receiver only Donald Trump’s or Bill Gates would have one. Fortunately there’s a way to get by with only moderately accurate clocks in our receivers — and the secret is to make an extra satellite range measurement. An extra distance measurement can make up for imperfect sync on our part. (Now you know why we were saying earlier that ”theoretically” three measurements are enough). Trigonometry says that if three perfect measurements locate a point in 3 D space, then four imperfect measurements can eliminate any timing offset (as long as the offset is consistent). The idea is really pretty simple. And it’s so fundamental to GPS. The explanation will be a lot easier to understand with diagrams, and those diagrams will be a lot easier to draw if we work in just two dimensions. Of course GPS is a three dimensional system, but the principle we’re discussing works the same in two dimensions. We just eliminate one measurement. Suppose our receiver’s clock isn’t perfect like an atomic clock. It’s consistent like a quartz watch but it’s not perfectly synced with universal time. OK, let’s say that, in reality, our distance from satellite S1 is R1 and R2 from satellite S2. In two dimensions, those two ranges would be enough to locate us at a point. Let’s call it ”X” (Remember, it takes three measurements to locate a point in three dimensions). So ”X” is where we really are and is the position we’d get if all the clocks were working perfectly. But now what if we used our ”imperfect” receiver? It would call the distance to satellite S1− R1 plus bias rB, and the distance to satellite S2− R2 plus bias rB. And that causes the two circles to intersect at a different point: ”XX”. HAM US/F-4 SaR Dec 2005 So XX is where our imperfect receiver would put us. And it would seem like a perfectly correct answer to us, since we’d have no way of knowing that our receiver was a little “imperfect”. But it would be miles off. We’d probably notice something wasn’t right when we started running into rocks, but nothing in the calculations would tell us. Now this is where our trigonometry trick can help: Let’s add another measurement to the calculation. In our two dimensional example, that means a third satellite S3. The ”pseudo−ranges” caused by our “imperfect “clock- the phrase ”pseudo− range” is used in GPS circles to describe ranges that contain errors (usually timing errors) will not result in one crossing. The result of the three LOPs ( LOP = Line of Position ) will be a triangle. The computers in our GPS receivers are programmed such that when they get a series of measurements that cannot intersect at a single point, they realize something is wrong. And they assume the cause is that their internal clock is off —that it has some offset. The computers now apply algebra to the problem. The old ”four equations and four unknowns” exercise. In three dimensions this means, that we really need to make four measurements to cancel out any error. And that’s a very significant number to remember because it means that you can’t get a truly accurate position until you have four satellites above the horizon around you. Summary S Accurate timing is key to measuring distance to satellites. S Satellites are accurate because they have atomic clocks on board. S Receiver clocks don’t have to be perfect because a trigonometry trick can cancel out receiver clock errors. S The trick is to make a fourth satellite range measurement. Page 444 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 X For Training Purposes Only XX Figure 227 HAM US/F-4 SaR Dec 2005 Pseudo Range Measuring Page 445 Part-66 SYSTEM OVERVIEW The GPS can be divided into three major groups: S Ground Control Segment S Space Segment S User Segment For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS HAM US/F-4 SaR Dec 2005 Page 446 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 228 HAM US/F-4 SaR Dec 2005 GPS System Segments Page 447 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 S Ground Control Segment GPS was developed by the US Department of Defense ( DoD ). The US Air Force has injected each satellite into a very precise orbit according to the GPS master plan. The orbits are known in advance and, in fact, some GPS receivers on the ground have an ”almanac” programmed into their computer’s memory, which tells them where in the sky each satellite will be at any given moment. Now this mathematical model of the orbits would be pretty accurate by itself, but, just to make things perfect, the GPS satellites are constantly monitored by the Department of Defense. That’s one of the reasons the GPS satellites are not put into geo−synchronous orbit like TV satellites are. The system is controlled by the Master Control Station in Colorado Springs and monitored by five Monitor Stations. The location of the monitor stations is choosen in such a way, so that every satellite can be seen from at least four stations at the same time during one day period. Since they go around the planet once every twelve hours, the GPS satellites pass over one of the DoD monitoring stations twice a day. This gives the DoD a chance to precisely measure their altitude, position and speed. The variations they’re looking for are called ”ephemeris” errors. They’re usually very minor and are caused by things like gravitational pulls from the moon and sun and by the pressure of solar radiation on the satellite. The master station is calculating the necessary parameters, describing the orbit and clock accuracy of each satellite. Once the DoD has measured a satellite’s position, they upload that information along with other data to the satellite. This computed orbits are called ephemeris. Necessary corrections are transmitted to the satellites via the ground antenna stations. The satellite will broadcast these minor corrections along with its timing information to the user. That’s an important fact to remember: S GPS satellites not only transmit a pseudorandom code for timing purposes but they also transmit a ”data message” about their exact orbital location and their system’s health. All serious GPS receivers use this information, along with the information in their internal almanacs, to precisely establish the position of the satellite. HAM US/F-4 SaR Dec 2005 Page 448 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 Master Control Station Ground Antenna Figure 229 HAM US/F-4 SaR Dec 2005 Backup Control Station Monitor Station GPS Operational Control Page 449 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 S Space Segment As already seen in the description of the GPS “ Pseudo Range “ measuring, four satellites ( Space Vehicle = SV ) are necessary for a 3D position. For a worldwide coverage 24 GPS SATELLITES are installed, 3 as active spares. There are 6 orbit planes with 4 satellites at each plane. The orbit height is about 20.000 km, so that it needs about 12 hours for one run. This constellation ensures at least 6 satellites in sight any time. Satellite’s Signal Each satellite transmits two signals for positioning purposes: S L1 centered on a carrier frequency of 1575.42 MHz S L2 centered on 1227.60 MHz. For Training Purposes Only L1 Signal Modulated onto the L1 carrier are two pseudo random noise (PRN) ranging codes: S the 1 mllisecond−long C/A−code with a chipping rate of about 1 megabit per second S a week−long segment of the normally encrypted P−code with a chipping rate of about 10 megahits per second. Also superimposed on the carrier is the navigation message which among other items includes the ephemeris data describing the satellite’s position and satellite clock correction terms. Civil receivers can only use this L1 signal and decode the C/A- code. Satellite’s position Because of the orbit hight of 20,200 km, the positions of the satellites is free of any earth influences and that means, that predictions of the satellites orbits will be very accurate. The orbits are known in advance and stored in a so−called ALMANAC. These almanac- data are transmitted within the navigation messages, so each receiver knows exactly where the satellites are at any time. Because it will take time to receive all the almanacs, modern receiver have an „ almanac“ programmed into their memory, which tells them, where in the sky each satellite will be at any given moment. Now this mathematical model of the orbits would be pretty accurate by itself, but, just to make things perfect, the GPS satellites are permanent monitored by the DoD. Any variations detected by the groundstation are called EPHEMERIS- errors. Once the DoD has measured a satellite’s position, they relay that information back up to the satellite. The satellites will broadcast these ephemeris data along with its timing information. All serious GPS receivers will use this ephemeris information together with their internal almanacs, to precisely establish the position of the SV. L2 Signal The L2 carrier is modulated only by the encrypted P−code and the navigation message. The C/A−code is not present. Because of national security concepts, the encrypted P−code (called the Y−code) is available only to authorized (primarily military) users through the Precise Positioning Service (PPS). The L2 signal is used primarily to reduce the effect of velocity changes given by the ionosphere. HAM US/F-4 SaR Dec 2005 Page 450 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 230 HAM US/F-4 SaR Dec 2005 21 GPS Satellites plus 3 Active Spares Page 451 Part-66 Satellites- Data The actual operating satellites are so-called Block 2 and 2A types manufactured by Rockwell Space Systems. Their designed lifetime is 7.5 years, the size with solar panel extended 5.3 m and the weight of one satellite is approx. 1.6 tons. Each SV carries two rubidium and two cesium clocks. The SV is 3− Axis stabalized with two solar arrays supplying 710 watts. The L-Band Antenna is a phased array with 12 helical antennas right hand circular polarized. It is used to downlink the frequencies L1 ( 1.575,42 MHz ) and L2 ( 1.227,6 MHz ). For Tracking, Telemetry & Control ( TT&C ) an S-Band antenna is used. The uplink is done on 1.783,74 MHz and downlink on 2.227,5 MHz. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS 28 Mar 96 16 July 96 12 Nov 96 HAM US/F-4 SaR Dec 2005 Page 452 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 231 HAM US/F-4 SaR Dec 2005 GPS Received Power Page 453 Part-66 S User Segment The GPS user segment in an aircraft consists of S antenna S receiver/processor unit S control and display unit The receiver convert the satellite sinals, received from the antenna, into the − position ( X,Y,Z ) − time − velocity GPS data, like: GPS position, Ground Speed, Status, Mode e.c. are displayed on GPS pages of the CDU. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS HAM US/F-4 SaR Dec 2005 Page 454 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 232 HAM US/F-4 SaR Dec 2005 GPS Major Components Page 455 Part-66 GPS Antenna There is a large number of GPS Antenna disigns. Modern types of L- Band antennas are top mounted, low profile active- or passivantennas. They operate at 1.575,42 MHz with right hand circular polarization ( RHCP ) and provide an omnidirectional upper hemispheric coverage so that signals within elevation angles from 10° to 90° can be received. The power for the active antennas is normally fed via the coaxialcable. For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS HAM US/F-4 SaR Dec 2005 Page 456 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 233 HAM US/F-4 SaR Dec 2005 GPS Antenna Page 457 Lufthansa Technical Training M13.4 NAVIGATION GPS Part-66 GPS Receivers The receivers can be divided into two maingroups: 1. C/A - Code - Receiver 2. P/Y - Code - Receiver For Training Purposes Only 1.) C/A - Code - Receiver The C/A ( Coarse Aquisition ) Code receiver is used by civil users. It can receive only the L1 frequency ( 1575.42 Mc ). The C/A Code (Coarse Acquisition) modulates the L1 carrier phase. The C/A code is a 1.023 MHz Pseudo Random Noise (PRN) Code. This noise−like code modulates the L1 carrier signal. The C/A code repeats every 1023 bits (one millisecond). There is a different C/A code PRN for each SV. GPS satellites are often identified by their PRN number, the unique identifier for each pseudo−random−noise code. The C/A code that modulates the L1 carrier is the basis for the civil SPS. The Navigation Message also modulates the L1−C/A code signal. The Navigation Message is a 50 Hz signal consisting of data bits that describe the GPS satellite orbits, clock corrections, and other system parameters. 2.) P/Y - Code - Receiver The P/Y Code receiver is used by military- or authorized- users only. It can receive the L1 and the L2 frequency ( 1227.60 Mc ). By receiving two frequencies, a receiver can correct changes of the propagation delay in the ionosphere.The L2 frequency is used to measure the ionospheric delay by PPS equipped receivers. The P−Code (Precise) modulates both the L1 and L2 carrier phases. The P−Code is a very long (seven days) 10.23 MHz PRN code. In the Anti−Spoofing (AS) mode of operation, the P−Code is encrypted into the Y−Code. The encrypted Y−Code requires a classified AS Module for each receiver channel and is for use only by authorized users with cryptographic keys. Only the use of the P (Y)−Code will give the high accuracy for the PPS. HAM US/F-4 SaR Dec 2005 C/A - Code Phase Tracking (Navigation) The GPS receiver produces replicas of the C/A−Code. Each PRN code is a noise− like, but predetermined, unique series of bits. The receiver produces the C/A code sequence for a specific SV in a C/A code generator. Modern receivers usually store a complete set of precomputed C/A code chips in memory. The C/A code generator repeats the PRN−code sequence every millisecond. The receiver slides the replica of the code in time until there is correlation with the SV code. If the receiver applies a different PRN code to an SV signal there is no correlation. When the receiver uses the same code as the SV and the codes begin to line up, some signal power is detected. As the SV and receiver codes line up completely, the full signal power is detected. A GPS receiver uses the detected signal power in the correlated signal to align the C/A code in the receiver with the code in the SV signal. Simplified GPS Receiver Block Diagram A phase locked loop that can lock to either a positive or negative half−cycle (a bi− phase lock loop) is used to demodulate the 50 HZ navigation message from the GPS carrier signal. The same loop can be used to measure and track the carrier frequency (Doppler shift) and by keeping track of the changes to the numerically controlled oscillator, carrier frequency phase can be tracked and measured. Data Bit Demodulation and C/A Code Control The receiver PRN code start position at the time of full correlation is the time of arrival (TOA) of the SV PRN at receiver. This TOA is a measure of the range to SV plus the amount, to which the receiver clock is offset from GPS time. This TOA is called the pseudo−range. Page 458 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION GPS Figure 234 HAM US/F-4 SaR Dec 2005 Simplified GPS Receiver Block Diagram Page 459 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 GPS SIGNAL STRUCTURE GPS- Frequencies The navigation signal transmitted from the space vehicles consists of two RF−frequencies: S L1 at 1575.42 MHz and S L2 at 1227.6 MHz. The L1 signal is modulated with both the C/A- and the P/Y- pseudo−random noise codes in phase quadrature. The L2 signal is modulated with the P/Y−code. Both the L1 and L2 signals are also continuously modulated with the navigation data−bit stream at 50 bps. All frequencies in the SV are first generated from the very stable atomic clocks and divided to the basicfrequency of fo = 10,23 MHz. All other frequencies are generated from this basic-frequency. Signal Basicfrequency Frequency 10,23 MHz fo Carrierfrequency L1−Carrier 1.575,42 MHz fo x 154 L2−Carrier 1.227,60 MHz fo x 120 For Training Purposes Only Codefrequency C/A−Code 1,023 MHz fo x 0,1 P/Y−Code 10,23 MHz fo GPS- Codes The functions of the codes are: 1. identification of space vehicles, as the code patterns are unique to each space vehicle and are matched with like codes generated in the user receiver 2. the measurement of the navigation signal transit time by measuring the phase shift required to match the codes. C/A- Code The C/A (clear access) code is a short code operating at 1.023 Mbps. The C/A code is normally acquired first and a transfer is made to the P−code by the use of the handover word (HOW) contained in the navigation data stream. The C/A code is a pseudo−random noise chip stream unique in pattern to each space vehicle that repeats every millisecond. It is relatively easy for the receiver to match and lock onto the C/A code. The total phase−shift required for lock−on is the measured pseudo−range time, including the offset in the user clock as well as the propagation delays and system errors. P/Y- Code The P ( Precise ) code is a long precision code operating at 10.23 Mbps but difficult to acquire. The P−code generated in each space vehicle is a pseudo−random noise chip sequence of seven days in length. That is, the pattern repeats only once every seven days. The P- code can be encrypted and is then called Y- code. This operation is named A-S ( anti spoofing ). Navigationfrequency Datamessage HAM US/F-4 SaR Dec 2005 50 bps Page 460 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 235 HAM US/F-4 SaR Dec 2005 GPS Signal Structure Page 461 Lufthansa Technical Training M13.4 NAVIGATION GPS Part-66 THE NAVIGATION MESSAGE The navigation message contains the data that the user’s receiver requires to perform the operations and computations for successful navigation with the GPS. The data include information on: S status of the space vehicle S time information for the transfer from the C/A to the P−code S parameters for computing the clock correction S corrections for delays in the signal propagation through the atmosphere S ephemeris of the space vehicle S almanac information of all space vehicles. For Training Purposes Only The navigation message is formatted in five subframes of six seconds in length, which make up a data frame of 30 seconds, 1500 bits long. The data are nonreturn−to−zero (NRZ) at 50 bps and are common to the P− and C/A signals on both the L1 and L2 channels. A space vehicle health−status word is provided in the fifth subframe (Data Block III) which indicates the status of the SV and permits the user the option of selecting another SV. The information in the navigation message is provided by the Control Segment and includes three blocks of data Block I to III. S Block I Block I data, which is contained in subframe 1, includes − clock correction parameters − approximation of the atmospheric delay for receivers using only the L1 signal and requiring less precision in their navigation − age of data (AOD) indicates the time since the last navigation upload for the use of more sophisticated receivers who may wish to select a satellite with a more recent update. HAM US/F-4 SaR Dec 2005 S Block II Block II data, which is contained in subframe 2 and 3, contains − ephemeris prediction parameters Both the clock correction parameters and the ephemeris parameters that are contained in Data Block II are updated every hour. S Block III Block III data, which is located in Subframe 4 and 5, contains − Almanac data These data include information on the ephemerides, clock correction parameters, and atmospheric delay parameters for the normal complement of 24 satellites, plus one spare. The data are a subset of Block I and Block Il parameters with reduced precision plus health and identification words for the space vehicles. The data are required to facilitate the rapid selection of space vehicles for use in the navigation solution. The receiver uses the information first to identify which satellites are in view and to solve the algorithm that indicates the SVs, that will provide the best navigation solution. The codes of these are then generated by the receiver for matching with the corresponding codes among all the incoming signals. The total Almanac data exceeds the capacity of single Subframes 4 and 5, so that it is transmitted on a rotating page basis. The complete Almanac is contained in 25 frames, so it will take a total of 12.5 minutes to transmit the Almanac data of all SVs. Sophisticated receivers will maintain almanacs in their data storage that preclude the need to wait for the transmission of the complete Almanac. Note: Almanacs are approximate orbital data parameters for all SVs. The ten−parameter almanacs describe SV orbits over extended periods of time (useful for months in some cases) and a set for all SVs is sent by each SV over a period of 12.5 minutes (at least). Signal acquisition time on receiver start−up can be significantly aided by the availability of current almanacs. The approximate orbital data is used to preset the receiver with the approximate position and carrier Doppler frequency (the frequency shift caused by the rate of change in range to the moving SV) of each SV in the constellation. Each complete SV data set includes an ionospheric model that is used in the receiver to approximates the phase delay through the ionosphere at any location and time. Page 462 Lufthansa Technical Training M13.4 NAVIGATION GPS Part-66 Block I Block II Block III S Ephemeris parameter ( actual satellite ) S Almanac for satellites 1−32 S Health of satellites 1−32 For Training Purposes Only S Clock correction S atmospheric delay correction S Age of data Figure 236 HAM US/F-4 SaR Dec 2005 GPS Data Message Page 463 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 GPS INTEGRITY How truthful is GPS? Can we believe the position that our GPS receiver computes? The GPS Standard Positioning Service is designed to provide a horizontal position accuracy of at least 100 meters, but such accuracy cannot be guaranteed 100 percent of the time. Satellite or ground system failures could cause a receiver to use erroneous data and compute positions that exceed its normal accuracy level. The performance of any navigation system is characterized by its S accuracy, S availability, S continuity, and S integrity. From a safety point of view, integrity is the most important factor. Without some assurance of a system’s integrity, we have no way of knowing whether the information we receive is correct. PERFORMANCE PARAMETERS S Accuracy. Accuracy describes how well a measured value agrees with a reference value. Ideally, the reference value should be the true value, if known, or some agreed− upon standard value.In terms of GPS positioning, a reference value might be the published coordinates of a geodetic reference mark. A measurement’s error is simply the difference between the obtained value and the reference value. GPS Accuracy The total GPS accuracy p is given by two factors: 1. Constellation or geometry of the used satellites ( DOP ) 2. Accuracy of the pseudo range measurement ( m ) p = GDOP x m For Training Purposes Only 1.) Dilution Of Precision ( DOP ) S GDOP means geometric dilution of precision (x,y,z,t) and depends on the relative position of the received and evaluated satellites. If the satellites bunches together, the failure in position is bigger (poor GDOP). The optimum position is one satellite overhead, 3 on the horizon, 120 apart in azimuth (ideal case, good GDOP). S PDOP is the position dilution of precision (x,y,z) S HDOP is the horizontal dilution of precision (x,y) S VDOP is the vertical dilution of precision (z) S TDOP is the time dilution of precision (t) HAM US/F-4 SaR Dec 2005 Page 464 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 237 HAM US/F-4 SaR Dec 2005 GPS Dilution Of Precision ( DOP ) Page 465 Lufthansa Technical Training M13.4 NAVIGATION GPS Part-66 2. ) Pseudo Range Measurement ( The total pseudo range error ( m ) m ) is given by a number of error sources. The User Equivalent Range Error ( UERE ) is calculated by the Root Sum Square ( RSS ). The largest error is produced by the ionospheric delay. Civil user can only minimize this error by using an ionospheric model. Military user with PPS can correct this error, because the delay changes with the frequency, so that with L1 and L2 the propagation delay can be corrected. For different frequencies the time delay is inversly proportional to the square of the carrier frequency: k T= f2 If only one frequency is used, a generic model of the ionospherics behavior (ionospheric modeling) may be used for a more precise position determination. R1 = Rtrue + k R1- f12 Rtrue = For Training Purposes Only R2 = Rtrue + k 1- f22 2 R f2 f1 2 2 f2 f1 R1 is the range to the satellite measured by use of f1. R2 respectively. Rtrue is the corrected range. This process is only used by PPS. SPS has only one frequency available: L1. HAM US/F-4 SaR Dec 2005 Page 466 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION GPS Figure 238 HAM US/F-4 SaR Dec 2005 GPS Error Sources Page 467 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION GPS Part-66 S Availability. A navigation system’s availability refers to its ability to provide the required function and performance within the specified coverage area at the start of an intended operation. In many cases, system availability implies signal availability, which is expressed as the percentage of time that the system’s transmitted signals are accessible for use. According to the 1996 Federal Radionavigation Plan, the GPS SPS will be available at least 99.85 percent of the time, based on a global average. S Continuity. Ideally, any navigation system should be continuously available to users. But, because of scheduled maintenance or unpredictable outages, a particular system may be unavailable at a certain time. Continuity is the ability of a total navigation system to function without interruption during an intended period of operation. S Integrity. The integrity of a navigation system refers to its honesty, veracity, and trustworthiness. A system might be available at the start of an operation, and we might predict its continuity at an advertised accuracy during the operation. But what if something unexpectedly goes wrong? If some system anomaly results in unacceptable navigation accuracy, the system should detect this and warn the user. Integrity characterizes a navigation system’s ability to provide this timely warning when it fails to meet its stated accuracy. Traditionally, some component of the navigation system or an independent monitoring unit assures integrity by examining the transmitted signals and providing a timely warning if they are out of specification. For example: VOR systems use an independent monitor to supply system integrity and remove a signal from use within 10 seconds of an out−of−tolerance condition. Integral monitors in ILS and MLS facilities similarly exclude anomalous signals from use within one second. S GPS INTEGRITY GPS achieves integrity and protects users against system anomalies and failures by relying on satellite self−checks and monitoring by the U.S. DoD (MCS), as well as signal assessment by users. Thus, GPS has both integral and independent mechanisms to assure integrity. S Satellite Self−Checks. GPS satellites internally monitor themselves for some, but not all, anomalies. These include navigation data errors, selective availability (SA) and antispoof HAM US/F-4 SaR Dec 2005 (AS) failures, and certain types of satellite clock failures. If such internal failures are detected, satellites notify users within six seconds. S Master Control Station. The GPS constellation is monitored by the MCS. Using data collected by five monitor stations distributed around the globe, MCS assesses GPS performance every 15 minutes by conducting tolerance and validation checks of the measured pseudoranges using a Kalman−filter, error−management process. In some circumstances, ranging errors could go undetected by this process for as long as 29 minutes. In addition to inform users of health problems, MCS issues Notice Advisories to Navstar Users (NANUs) that report satellite outages as well as planned service losses caused by maintenance. AVIATION REQUIREMENTS RAIM is a necessary component for GPS aviation applications. The U.S.FAA has mandated that any GPS receiver used for supplemental aircraft navigation incorporate integrity monitoring. Table 1 ( Figure 239 ) lists the FAA integrity performance requirements for the various phases of flight. Currently, with enhancements to RAIM, aviators can even use GPS as a primary−means navigation system for the en route oceanic phase of flight as well as in remote areas (see Table 2 for navigation requirements of different flight phases). Employing GPS for sole−means navigation during other phases of flight will require augmentation systems to reduce the probability of attaining misleading information within a one−hour period to 10−7. Aircraft have different needs depending on where they are. Over the North Atlantic e.g., aircraft must maintain S a cross−track position accuracy of 12.6 nautical miles, 2 sigma ( 95% ), and S a height accuracy of 350 feet, 3 sigma ( 99.7% ) . To conduct a Category III precision approach to a runway, S the cross−track accuracy at the runway threshold must be better than 4.1 meters (95 percent), S with vertical accuracy better than 0.6 meters (95 percent). Page 468 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION GPS Figure 239 HAM US/F-4 SaR Dec 2005 GPS Integrity / Navigation Requirements Page 469 Part-66 RAIM Although the GPS command segment and the satellites themselves provide a reasonable level of integrity for most GPS purposes, it is insufficient for aviation, as anomalies could go undetected for too long a period. It typically takes MCS 5−15 minutes to remove a satellite with a detected anomaly from service. Furthermore, if a satellite is not in view of one of the ground stations an anomaly could go undetected for more than 10 minutes. The need for a higher level of integrity for GPS use in aviation led to the concept of having a GPS receiver independently or autonomously establish system integrity. This concept is known as receiver autonomous integrity monitoring, or RAIM. RAIM attempts to answer two questions: S Is there a GPS satellite failure? S Which is the errant satellite? If GPS is used solely for supplemental navigation, then answering the first question is sufficient, because an alternative navigation system is available. However, if GPS is used for primary−means navigation, both questions must be answered to identify and remove the failed satellite from the solution, allowing the aircraft to safely proceed. Answering either question necessitates measurement redundancy; that is, more than the minimum four measurements required for a position solution. We need measurements from at least five satellites to detect a satellite anomaly and a minimum of six to remove the faulty satellite from the navigation solution. RAIM AVAILABILITY One possible way to determine satisfactory satellite geometry involves adopting a maximum permissible horizontal dilution of precision (HDOP) value and declaring all geometries with HDOP greater than HDOPmax inadmissible. An aircraft GPS receiver must have at least six satellites in view above an elevation mask angle of 7.5 degrees to provide RAIM ( detect and remove ). This condition is not always satisfied with the existing GPS constellation, and in these situations, RAIM would be unavailable. Redundancy can be increased by adding measurements from other systems, however, such as GLONASS and onboard inertial reference units (IRU) or barometric altitude measurements, which effectively provide an additional range value. With such auxiliary measurements, RAIM may be available where, with GPS measurements alone, it would not be. For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION GPS HAM US/F-4 SaR Dec 2005 Page 470 Part-66 THIS PAGE INTENTIONALLY LEFT BLANK For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION GPS HAM US/F-4 SaR Dec 2005 Page 471 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 GPS NAVIGATION SERVICES Three services are available from GPS: S Standard Positioning Service (SPS) for civil and military users: − Single frequency (L1) operation − Low rate PN sequence used (C/A code at 1.023 Mbps) − Degraded accuracy for non-DoD (Department of Defence) users (selective availability) S horizontal 100m S vertical 150m S 3D position 120m S time 350nsec S velocity vector 0.3m − Undegraded accuracy for autorized users: 20-40m − Civil accuracy degradable when US-national policy dictates. S Precise Positioniong Service (PPS) for military users with more stringent requirements − Dual frequency (L1,L2) operation − High rate PN sequence used (P or Y code at 10.23 Mbps) − Full system accuracy for authorized users S 3D position 16m S 2D position 17.8m S altitude 27.7m S time 200nsec S velocity vector 0.1m/sec − Other characteristics of PPS: S Anti-jam capabilities S Anti-spoof encryption op P code (Y) S Selective availability (SA) degradation − Uses decryption techniques to acess the capabilities S Differential positioning GPS (code) − 1-5m 2DRMS HAM US/F-4 SaR Dec 2005 DIFFERENTIAL GPS Measurement precision can be improved by using differential methods. In the area of the satellite navigation system a second receiver is exactly positioned. Errors occuring in the system (time bias due to ionospheric-, tropospheric effects, etc.) are determined by this receiver, because its location is very well known. This information is transmitted to the user receiver, which is able now to determine its own position with higer accuracy. REFERENCE STATION Page 472 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 Normal Operation ( PPS ) Differential Operation For Training Purposes Only Error Sources Navigationaccuracy Error Sources GDOP x UERE 2,3 x 7,0 m 16 m Figure 240 HAM US/F-4 SaR Dec 2005 Navigationaccuracy GDOP x UERE 2,3 x 1,3 m 3m GPS / DGPS Navigationaccuracy Page 473 Lufthansa Technical Training For Training Purposes Only M13.4 NAVIGATION GPS Part-66 GPS OPERATING MODES The GPS Receiver (GPSSU) has seven operating modes : S Self−Test, Initialization, Acquisition, Navigation, Altitude/Clock Aiding, Aided andFault. The device transitions between modes automatically. The current GPSSU operating mode can be known by selecting the GPS MONITOR page on the FMS CDU. In Navigation mode, valid input data can be used to generate the navigation solution. S Self−Test mode The GPSSU is in Self−Test mode during the period from the power application until completion of all internal power−up BITE. The time duration spent in Self− Test mode is no more than 5 seconds. S Initialization mode The GPSSU is in Initialization mode during the period of time in which it has completed Self−Test mode until the device has initialized the hardware to enable it to enter Acquisition mode. S Acquisition mode The GPSSU is in Acquisition mode when insufficient satellite and/or aiding data are available to produce an initial navigation solution. To acquire signals from the GPS satellites, the GPSSU uses : −almanac data which describe the satellite orbits, −time which, in conjunction with almanac data, is used to estimate the present position of satellites in their orbits. −the appropriate location of the GPSSU so a prediction can be made as to which satellites are visible. The GPSSU stores almanac data in nonvolatile memory which does not require an internal or external battery for support. The GPSSU predicts which satellites are visible and acquires the satellite signals. The GPSSU then collects ephemeris data by decoding the satellite down−link data message. After each satellite in view is acquired, the satellite measurement data are transmitted continuously. When a sufficient number of satellites are being tracked, position and velocity can be computed and Navigation mode entered. HAM US/F-4 SaR Dec 2005 S S S S −Search the Skies If the GPSSU cannot perform an acquisition, due to an absence of almanac data, it initiates a ”Search the Skies” acquisition. The GPSSU attempts to acquire all visible satellites in the GPS constellation. Once a satellite has been acquired, ephemeris data are decoded from the satellite downlink message. After sufficient satellites have been acquired, the GPSSU sets the SSM to the appropriate status mode and enters Navigation mode. Without valid initialization, the Time To First Fix (TTFF) is not more than 600 seconds. Navigation mode The GPSSU enters in Navigation mode when the GPSSU has decoded ephemeris data to compute a navigation solution, and when the number of GPS satellites being tracked provides a sufficient set of measurements to compute position, velocity, and time outputs. In this mode, the satellite measurement data continue to be transmitted without interruption. Altitude/Clock Aiding mode If satellite measurements are not sufficient to perform in Navigation mode, yet are sufficient when altitude and clock information is available, it is in Altitude/ Clock Aiding mode. This mode uses inertial altitude and clock drift information to aid the navigation solution and integrity monitoring during extended periods of insufficient satellite coverage and geometry. The GPSSU enters Altitude/ Clock Aiding mode only after the pressure altitude has been calibrated with a geometric altitude solution. Aided mode If satellite measurements are not sufficient to perform in Navigation mode or Altitude/Clock Aiding mode, but external aiding data is available, the GPSSU could be in Aided mode. This mode may use inertial velocities to aid the navigation solution and integrity monitoring during extended periods of insufficient satellite coverage and geometry. The GPSSU may enter Aided mode only after it has been determined that there is not sufficient satellite and calibrated altitude data available to remain in Altitude/Clock Aiding mode or in Navigation mode. In this mode, all the parameters are NCD . It allows to maintain the receiver position using external data (e.g. IRS Lat/Long,...) in order to perform a quicker acquisition of the satellites necessary to re−enter in Navigation mode. Fault mode The GPSSU is in Fault mode during the period of time in which the device outputs are affected by one or more severe system faults. This mode supersedes all others and remains active until the next power−down/power−up cycle. Fault mode is entered from any other mode. Page 474 Part-66 For Training Purposes Only Lufthansa Technical Training M13.4 NAVIGATION GPS Figure 241 HAM US/F-4 SaR Dec 2005 GPS Operating Modes Page 475 Lufthansa Technical Training For Training Purposes Only M11.5.2 / M13.4 NAVIGATION GPS Part-66 DISPLAY OF GPS DATA ON FMS CDU POSITION MONITOR page (FIGURE 242 ) GPS primary navigation function in the FMCs A navigation mode with the least error is chosen based upon the mixed IRS position and the best GPS or radio position available. NOTE:The GPS position used by the FMCs to determine the aircraft position is computed in the GPSSU. The mode of navigation is selected according to the following hierarchy : −GPS/Inertial −DME/DME/Inertial −DME/VOR/Inertial−Inertial only. The GPS/INERTIAL mode is selected as long as the following conditions are satisfied : −GPS position available and with an estimated accuracy consistent with the intended operation. −GPS integrity (provided by the RAIM) is available and compatible with the requirement for the applicable phase of flight. The N IRS/GPS indication is displayed on the POSITION MONITOR page on the FMS CDU with N being the number of IRS used to compute mixed IRS position. The selected GPS position is displayed on the POSITION MONITOR page in place of radio position. Aircraft position is generated by a series of filters which use inertial position, GPS position or radio position, and aircraft velocity as inputs. The GPS/INERTIAL mode can be manually inhibited by pushing the line key adjacent to the DESELECT GPS indication on the POSITION MONITOR page on the FMS CDU. FMC computed integrity (AIM). When the GPS position is available in the FMC but the GPS integrity is not delivered by the GPSSU, the FMC is capable of computing an equivalent integrity called AIM (Alternate Integrity Monitoring) using IRS data during a limited period of time. HAM US/F-4 SaR Dec 2005 GPS MONITOR page The GPS data are displayed on the GPS MONITOR page of the FMS CDU. The upper part is dedicated to GPS 1 data, the lower part to GPS 2 data. The displayed data are : − GPS position (lat/long) − true track − GPS altitude − figure of merit (in meters) − ground speed − number of satellites tracked − mode. Page 476 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 242 HAM US/F-4 SaR Dec 2005 GPS Monitor Page Page 477 Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Part-66 PREDICTIVE GPS page (Figure 243 Part A) The integrity prediction results done by the GPSSU on FMS request are displayed on the PREDICTIVE GPS page of the CDU (from the REF page). The prediction concerns the destination (DEST) and any pilot entered waypoint (WPT) and the integrity availability (HIL < 0.3 Nm) is displayed by Yes (Y) or No (N) for the seven times defined by the five− minute increments for plus or minus 15 minutes around DEST or WPT. Display of GPS PRIMARY LOST message on NDs (Figure 243 ) When the GPS mode is active, the MSG indication is displayed on the ND and the GPS PRIMARY indication is displayed on the Progress page on the FMS CDU. When the GPS mode is not active, the MSG indication is displayed on the ND, with the GPS PRIMARY LOST message permanently displayed below. The GPS PRIMARY LOST indication is displayed on the Progress page on the FMS For Training Purposes Only Progress page (Figure 243 Part B) The progress page indicates whether the GPS is used by the FMCs for navigation. If the GPS is used by the FMCs for navigation, the GPS PRIMARY indication is displayed. If it is not used, the GPS PRIMARY LOST message is shown. Line 5 indicates: S on the left: − the Required Navigation Performance RNP ( for this flight segment it is entered 1.0 NM) S on the right: − the Estimated Position Error EPE ( e.a. 0.16 NM ) S in the center: − the Accuracy S HIGH: RNP > EPE S LOW: EPE > RNP HAM US/F-4 SaR Dec 2005 Page 478 Part-66 For Training Purposes Only Lufthansa Technical Training M11.5.2 / M13.4 NAVIGATION GPS Figure 243 HAM US/F-4 SaR Dec 2005 GPS Predictive Page Page 479 P66 B2 M13 4 NAV E TABLE OF CONTENTS ATA 34 NAVIGATION . . . . . . . . . . . . . . . . . . . . . . . . . ABBREVIATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2 VOR 12 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BEARING DEFINITION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR PRINCIPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR GROUNDSTATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR RANGE AND ACCURACY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR AIRCRAFT EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOMATIC VOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MANUAL VOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR ANTENNA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR RECEIVER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR TUNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR DISPLAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 16 18 20 28 30 32 34 38 40 44 50 ILS . 54 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS PRINCIPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS GROUND FACILITIES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOCALIZER THEORY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GLIDESLOPE THEORY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS AIRCRAFT EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS ANTENNAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS RECEIVER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS TUNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS DISPLAYS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54 56 58 60 66 72 74 76 78 82 MARKER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86 M L S ..... 92 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92 GROUND STATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 96 MLS PRINCIPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98 AIRCRAFT EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 AUTOMATIC DIRECTION FINDER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 104 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GROUND STATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF PRINCIPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF CONTROL PANEL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF INDICATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENTS LOCATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF RECEIVER SCHEMATIC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . APPENDIX . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 106 108 116 118 122 126 128 132 EGPWS .................................................... GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS EVOLUTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENVELOPE MODULATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TERRAIN AWARENESS ALERTING AND DISPLAY . . . . . . . . . . . . . . TERRAIN ALERTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TERRAIN DISPLAYS AND ALERTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . TERRAIN BACKGROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TERRAIN CLEARANCE FLOOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TERRAIN-WEATHER INTERFACE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MKV EGPWS INTERFACE CONTROL PROGRAM PINS . . . . . . . . . . MKV EGPWS PROGRAMMING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS: SELF TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS: STATUS LEDS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142 142 144 146 162 164 166 168 170 174 176 178 182 184 190 RADAR FUNDAMENTALS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RANGE DETERMINATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RADAR SYSTEM PARAMETERS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RADAR SOLUTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WAVE GUIDES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MICROWAVE GENERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SEMICONDUCTOR MICROWAVE DEVICES . . . . . . . . . . . . . . . . . . . . FREQUENCY DETERMINING ELEMENTS . . . . . . . . . . . . . . . . . . . . . . RADAR ANTENNAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194 194 196 198 204 208 220 226 234 238 Page i P66 B2 M13 4 NAV E TABLE OF CONTENTS DIGITAL INTERFACES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENT DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MEASUREMENT OF INTRUDER PARAMETERS . . . . . . . . . . . . . . . . DEFINITION OF TARGET AIRCRAFT . . . . . . . . . . . . . . . . . . . . . . . . . . ADVISORY INHIBIT CONDITIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SENSITIVITY LEVEL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INFORMATION DISPLAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SELF TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 340 342 348 356 364 368 370 372 382 288 288 290 296 300 310 312 318 320 326 WEATHER RADAR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRINCIPLE OF OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT EQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENTS LOCATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX TRANSCEIVER (HIGH POWER) . . . . . . . . . . . . . . . . . . . . . . . . . . . WX TRANSCEIVER (LOW POWER) . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX CONTROL FUNCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX−RADAR OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX ANTENNA STABILIZATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INDICATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SAFETY PRECAUTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TURBULENCE / WINDSHEAR GENERAL . . . . . . . . . . . . . . . . . . . . . . . DETECTION OF PRECIPITATION TARGETS . . . . . . . . . . . . . . . . . . . . PULSE PROPAGATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TURBULENCE DETECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WINDSHEAR DETECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WINDSHEAR ACTIVATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WINDSHEAR ALERT REGIONS AND LEVEL . . . . . . . . . . . . . . . . . . . . WINDSHEAR INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WXR / WINDSHEAR - TIME SHARING . . . . . . . . . . . . . . . . . . . . . . . . . 384 384 384 386 388 390 392 402 406 408 412 416 418 420 422 424 426 430 430 432 434 TCAS . 328 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 328 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 330 GENERAL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 336 GLOBAL POSITIONING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ABBREVIATIONSLIST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SATELLITE RANGING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 436 436 438 440 DME DME DME DME DME DME DME DME DME DME 244 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRINCIPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHANNEL-FREQUENCY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DATA DISPLAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MODES OF OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CONTROL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SCANNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ANTENNA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 244 246 248 252 254 256 258 258 260 LRRA 262 LRRA GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA GENERAL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA PRINCIPLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA SYSTEM FAILURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA DATA DISPLAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA OUTPUTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA ANTENNAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA TESTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 262 264 266 274 276 278 280 282 284 ATC TRANSPONDER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRINCIPLE OF OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATCRBS INTERROGATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MODE S INTERROGATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATCRBS REPLIES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MODE S REPLIES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC TYPICAL LOCATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENT DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC MODE S TRANSPONDER SELFTEST . . . . . . . . . . . . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 Page ii P66 B2 M13 4 NAV E TABLE OF CONTENTS DISTANCE MEASURING FROM A SATELLITE . . . . . . . . . . . . . . . . . . PSEUDO RANGE MEASURING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SYSTEM OVERVIEW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS SIGNAL STRUCTURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS INTEGRITY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS NAVIGATION SERVICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DIFFERENTIAL GPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS OPERATING MODES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DISPLAY OF GPS DATA ON FMS CDU . . . . . . . . . . . . . . . . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 442 444 446 460 464 472 472 474 476 Page iii P66 B2 M13 4 NAV E TABLE OF CONTENTS HAM US/F-4 Sabrotzki Dec 2005 Page iv P66 B2 M13 4 NAV E TABLE OF FIGURES Figure 1 Figure 2 Figure 3 Figure 4 Figure 5 Figure 6 Figure 7 Figure 8 Figure 9 Figure 10 Figure 11 Figure 12 Figure 13 Figure 14 Figure 15 Figure 16 Figure 17 Figure 18 Figure 19 Figure 20 Figure 21 Figure 22 Figure 23 Figure 24 Figure 25 Figure 26 Figure 27 Figure 28 Figure 29 Figure 30 Figure 31 Figure 32 Figure 33 Figure 34 Figure 35 VOR General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR Approach Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR Arrival ( STAR ) Chart ( Jeppesen ) . . . . . . . . . . . . . Magnetic- / Relative- Bearing . . . . . . . . . . . . . . . . . . . . . . . Lighthouse principle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR conventional ground station . . . . . . . . . . . . . . . . . . . . Conventional VOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conventional VOR Spectrum . . . . . . . . . . . . . . . . . . . . . . . DVOR ground station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DVOR Antenna installation . . . . . . . . . . . . . . . . . . . . . . . . DVOR Spectrum . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR- Range / - Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . VOR System Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR Automatic Indications . . . . . . . . . . . . . . . . . . . . . . . . VOR Deviation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR TO / FROM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR manual Indications . . . . . . . . . . . . . . . . . . . . . . . . . . VOR Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR Receiver block diagram . . . . . . . . . . . . . . . . . . . . . . VOR Termination loads . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR tuning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR Back Up tuning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR ND Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VOR RDDMI Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS Approach Chart ( FRA ) . . . . . . . . . . . . . . . . . . . . . . . ILS Principle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS Ground Facilities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOC Antenna Array ( HAM Rwy 23 ) . . . . . . . . . . . . . . . . LOC DDM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOC Range / DDM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOC Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GS Antenna Array . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GS DDM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GS Range / DDM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GS Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 13 14 15 17 19 20 21 23 24 25 27 29 31 33 35 36 37 39 41 43 47 49 51 53 55 57 59 60 61 63 65 66 67 69 71 Figure 36 Figure 37 Figure 38 Figure 39 Figure 40 Figure 41 Figure 42 Figure 43 Figure 44 Figure 45 Figure 46 Figure 47 Figure 48 Figure 49 Figure 50 Figure 51 Figure 52 Figure 53 Figure 54 Figure 55 Figure 56 Figure 57 Figure 58 Figure 59 Figure 60 Figure 61 Figure 62 Figure 63 Figure 64 Figure 65 Figure 66 Figure 67 Figure 68 Figure 69 Figure 70 ILS System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS Antennas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS Standard Deflection . . . . . . . . . . . . . . . . . . . . . . . . . . . ILS tuning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOC Indication front beam . . . . . . . . . . . . . . . . . . . . . . . . . G/S Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Marker System Schematic . . . . . . . . . . . . . . . . . . . . . . . . Marker Ground Station . . . . . . . . . . . . . . . . . . . . . . . . . . . . MKR Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Marker Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MLS and ILS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MLS - ILS Azimuth / Elevation . . . . . . . . . . . . . . . . . . . . . MLS Ground Stations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MLS Measurement Principle . . . . . . . . . . . . . . . . . . . . . . . MLS Transmission Format . . . . . . . . . . . . . . . . . . . . . . . . MLS Approach Capabilities . . . . . . . . . . . . . . . . . . . . . . . . ADF - NDB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Locators on Approach Maps . . . . . . . . . . . . . . . . . . . . . . NDBs in Jeppesen Maps . . . . . . . . . . . . . . . . . . . . . . . . . . ADF Loop Antenna Function . . . . . . . . . . . . . . . . . . . . . . ADF Ambiguity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF-Goniometer principle . . . . . . . . . . . . . . . . . . . . . . . . . Quadrantal Error . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF Blockdiagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ident and Voice Frequency Range . . . . . . . . . . . . . . . . . . Relative Bearing RB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF RMIs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF Components Locations . . . . . . . . . . . . . . . . . . . . . . . ADF Receiver . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF Receiver . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compass Information for the RMIs . . . . . . . . . . . . . . . . . EFIS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ND - ADF Pointers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADF - Data and Warning Displays . . . . . . . . . . . . . . . . . . 73 75 77 81 83 85 87 89 90 91 93 95 97 99 101 103 105 106 107 109 111 113 115 117 119 121 123 125 127 129 131 133 135 136 137 Page i P66 B2 M13 4 NAV E TABLE OF FIGURES Figure 71 Figure 72 Figure 73 Figure 74 Figure 75 Figure 76 Figure 77 Figure 78 Figure 79 Figure 80 Figure 81 Figure 82 Figure 83 Figure 84 Figure 85 Figure 86 Figure 87 Figure 88 Figure 89 Figure 90 Figure 91 Figure 92 Figure 93 Figure 94 Figure 95 Figure 96 Figure 97 Figure 98 Figure 99 Figure 100 Figure 101 Figure 102 Figure 103 Figure 104 Figure 105 Navigation Frequency Tuning . . . . . . . . . . . . . . . . . . . . . . Radio Navigation Back-Up . . . . . . . . . . . . . . . . . . . . . . . . . Type of Accident . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aircraft Accidents Statistic . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS Simplified Schematic . . . . . . . . . . . . . . . . . . . . . Mode 1 Excessive Descent Rate . . . . . . . . . . . . . . . . . . . Mode 2 Excessive Terrain Closure Rate . . . . . . . . . . . . . Mode 3 Altitude Loss After Takeoff . . . . . . . . . . . . . . . . . Mode 4 Unsafe Terrain Clearance . . . . . . . . . . . . . . . . . . Mode 5 Excessive Glide Slope Deviation . . . . . . . . . . . . Mode 6 Call-Outs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mode 7 Excessive Windshear Detection ( RWS ) . . . . . Envelope Modulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Terrain Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Terrain Alerting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TERRAIN INDICATION ON ND . . . . . . . . . . . . . . . . . . . . Terrain Background Display . . . . . . . . . . . . . . . . . . . . . . . . Terrain Clearance Floor . . . . . . . . . . . . . . . . . . . . . . . . . . . Terrain / WXR Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aircraft Type Strapping Table D-T1 . . . . . . . . . . . . . . . . . Basic Audio Menu . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS: SELF TEST, LEVEL 1 . . . . . . . . . . . . . . . . . . . . EGPWS: SELF TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS: SELF TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGPWS: COMPUTER STATUS LEDs . . . . . . . . . . . . . . Radar Range Determination . . . . . . . . . . . . . . . . . . . . . . . Radar Transmit- / Echo- Power . . . . . . . . . . . . . . . . . . . . Radar Attenuation and Reflectivity of E.M.Waves . . . . . Radar PRF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radar Solution Criterion: Pulse Width . . . . . . . . . . . . . . Radar Solution Criterion: Beam Width . . . . . . . . . . . . . . Radar Coaxial Cable and Wave Guide . . . . . . . . . . . . . Radar Modes in Wave guides . . . . . . . . . . . . . . . . . . . . . Radar Wave Propagation in Wave Guides . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 139 141 143 145 147 149 151 153 155 157 159 161 163 165 167 169 173 175 177 179 181 183 185 187 189 191 197 199 201 203 205 207 209 211 213 Figure 106 Figure 107 Figure 108 Figure 109 Figure 110 Figure 111 Figure 112 Figure 113 Figure 114 Figure 115 Figure 116 Figure 117 Figure 118 Figure 119 Figure 120 Figure 121 Figure 122 Figure 123 Figure 124 Figure 125 Figure 126 Figure 127 Figure 128 Figure 129 Figure 130 Figure 131 Figure 132 Figure 133 Figure 134 Figure 135 Figure 136 Figure 137 Figure 138 Figure 139 Figure 140 Radar Velocities in Wave Guides . . . . . . . . . . . . . . . . . . Radar Wave Guide Junctions . . . . . . . . . . . . . . . . . . . . . Radar Coupling methodes . . . . . . . . . . . . . . . . . . . . . . . . Radar Magnetron . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radar Klystron . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radar Reflex Klystron . . . . . . . . . . . . . . . . . . . . . . . . . . . . Microwave Semiconductors . . . . . . . . . . . . . . . . . . . . . . . Radar Gunn Diode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radar IMPATT Diode . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radar Tunnel Diode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Two Wires Resonator . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tank Circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cavity Resonator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Parabolic Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flat Plate Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Parabolic and Flat Plate Antenna . . . . . . . . . . . . . . . . . . DME General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DME Principle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DME Channel-Frequency . . . . . . . . . . . . . . . . . . . . . . . . DME Channel-Frequency Chart # 1 . . . . . . . . . . . . . . . . DME Channel-Frequency Chart # 2 . . . . . . . . . . . . . . . . DME Syst. Schematic ( A 320 ) . . . . . . . . . . . . . . . . . . . DME Data Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DME Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DME Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DME Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA Gweneral . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA General System Schematic . . . . . . . . . . . . . . . . . LRRA Pulse principle . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA with Constant Modulation Period . . . . . . . . . . . . LRRA with Constant Difference Frequency . . . . . . . . . Tracking height less than 5000 ft . . . . . . . . . . . . . . . . . . LRRA System Failures . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA Syst. Schematic ( A 320 ) . . . . . . . . . . . . . . . . . . LRRA Indicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 217 219 221 223 225 227 229 231 233 235 236 237 239 241 243 245 247 249 250 251 253 255 257 259 261 263 265 267 269 271 273 275 277 279 Page ii P66 B2 M13 4 NAV E TABLE OF FIGURES Figure 141 Figure 142 Figure 143 Figure 144 Figure 145 Figure 146 Figure 147 Figure 148 Figure 149 Figure 150 Figure 151 Figure 152 Figure 153 Figure 154 Figure 155 Figure 156 Figure 157 Figure 158 Figure 159 Figure 160 Figure 161 Figure 162 Figure 163 Figure 164 Figure 165 Figure 166 Figure 167 Figure 168 Figure 169 Figure 170 Figure 171 Figure 172 Figure 173 Figure 174 Figure 175 LRRA Analog Output Standard . . . . . . . . . . . . . . . . . . . LRRA Antennas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRRA Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Traffic Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC Ground Radar Scope . . . . . . . . . . . . . . . . . . . . . . . ATC Interrogation Principle . . . . . . . . . . . . . . . . . . . . . . . ATCRBS Interrogation Codes . . . . . . . . . . . . . . . . . . . . . ATCRBS Side Lobe Suppression . . . . . . . . . . . . . . . . . . All Call Interrogation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mode S Interrogation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Differential Phase Shift Keying (DPSK) . . . . . . . . . . . . . Mode S Side Lobe Suppression . . . . . . . . . . . . . . . . . . . Uplink Formats . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATCRBS Replies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mode S Replies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Downlink Formats . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . ATC Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC Transponder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC / TCAS Cont. PNL . . . . . . . . . . . . . . . . . . . . . . . . . . ATC Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ATC CFDS Menu . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS-Surveillance Envelope . . . . . . . . . . . . . . . . . . . . . TCAS Target Classification . . . . . . . . . . . . . . . . . . . . . . . TCAS encounter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS Coordination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS General System Schematic . . . . . . . . . . . . . . . . . TCAS − PWR, A/ D INTERFACES AND PPs . . . . . . . TCAS − DIGITAL INTERFACES . . . . . . . . . . . . . . . . . . TCAS Processor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS / ATC Cont. PNL . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS ANTENNA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 281 283 285 286 287 291 293 295 297 299 301 303 305 307 309 311 313 315 317 318 319 321 323 325 327 329 331 333 335 337 339 341 343 345 347 Figure 176 Figure 177 Figure 178 Figure 179 Figure 180 Figure 181 Figure 182 Figure 183 Figure 184 Figure 185 Figure 186 Figure 187 Figure 188 Figure 189 Figure 190 Figure 191 Figure 192 Figure 193 Figure 194 Figure 195 Figure 196 Figure 197 Figure 198 Figure 199 Figure 200 Figure 201 Figure 202 Figure 203 Figure 204 Figure 205 Figure 206 Figure 207 Figure 208 Figure 209 Figure 210 TCAS Whisper-Shout . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transponder Mode A/C Reply Format . . . . . . . . . . . . . . TCAS -Mode S-Interrogation and Reply . . . . . . . . . . . . TCAS used UP- / Downlink Formats . . . . . . . . . . . . . . . TCAS Intruder Parameters . . . . . . . . . . . . . . . . . . . . . . . TCAS Coordination DF11 / UF0 / DF0 . . . . . . . . . . . . . TCAS Coordination UF16 / DF16 . . . . . . . . . . . . . . . . . . TCAS Vertical Separation / TAU . . . . . . . . . . . . . . . . . . . TCAS Corrective / Preventive Advisory Thresholds . . TCAS Airborn / Ground Definition . . . . . . . . . . . . . . . . . TCAS Senitivity Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS Indication on ND . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS ROSE- / ARC- Mode Display . . . . . . . . . . . . . . . TCAS Indication on PFD . . . . . . . . . . . . . . . . . . . . . . . . . TCAS PFD Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . TCAS Messages on ND/PFD/EWD . . . . . . . . . . . . . . . . TCAS - Test Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WEATHER RADAR SYSTEM . . . . . . . . . . . . . . . . . . . . Aircraft Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Components Locations . . . . . . . . . . . . . . . . . . . . . . . . . . WX-System Block Diagram . . . . . . . . . . . . . . . . . . . . . . . WX Radar Transceiver Blockdiagram . . . . . . . . . . . . . . WX Radar Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . WX-STC / Penetration Compensation . . . . . . . . . . . . . . WX Contour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX Video Digitizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX Video digitizing-Schematic . . . . . . . . . . . . . . . . . . . . ARINC 429 Control Word . . . . . . . . . . . . . . . . . . . . . . . . ARINC 453 WX-Data Word . . . . . . . . . . . . . . . . . . . . . . . PPI and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Antenna Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . . Single Axis Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . Split Axes Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . Plan Position Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . Weather-/Turbulence- Mode . . . . . . . . . . . . . . . . . . . . . . 349 351 353 355 357 359 361 363 365 369 371 373 375 377 379 381 383 385 387 389 390 391 393 395 397 399 401 403 405 407 409 410 411 413 414 Page iii P66 B2 M13 4 NAV E TABLE OF FIGURES Figure 211 Figure 212 Figure 213 Figure 214 Figure 215 Figure 216 Figure 217 Figure 218 Figure 219 Figure 220 Figure 221 Figure 222 Figure 223 Figure 224 Figure 225 Figure 226 Figure 227 Figure 228 Figure 229 Figure 230 Figure 231 Figure 232 Figure 233 Figure 234 Figure 235 Figure 236 Figure 237 Figure 238 Figure 239 Figure 240 Figure 241 Figure 242 Figure 243 Map Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety Distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wetter Detection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Weather Return Signal . . . . . . . . . . . . . . . . . . . . . . . . . . . Transmit Pulse Propagation . . . . . . . . . . . . . . . . . . . . . . Frequency spectrum . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Return Signal with Target Spectrum . . . . . . . . . . . . . . . Approaching a Microburst during Landing . . . . . . . . . . Determining Windshear Hazard . . . . . . . . . . . . . . . . . . . Detecting Microburst with Doppler Shift . . . . . . . . . . . . Windshear Alert Regions . . . . . . . . . . . . . . . . . . . . . . . . . Windshear Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . WXR - / FLW - Operation Time Sharing . . . . . . . . . . . . Comparison of Navigation Accuracy . . . . . . . . . . . . . . . GPS Satellite Ranging . . . . . . . . . . . . . . . . . . . . . . . . . . . Acquisition of the GPS C/A Signal . . . . . . . . . . . . . . . . . Pseudo Range Measuring . . . . . . . . . . . . . . . . . . . . . . . . GPS System Segments . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Operational Control . . . . . . . . . . . . . . . . . . . . . . . . . 21 GPS Satellites plus 3 Active Spares . . . . . . . . . . . . GPS Received Power . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Major Components . . . . . . . . . . . . . . . . . . . . . . . . . GPS Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Simplified GPS Receiver Block Diagram . . . . . . . . . . . GPS Signal Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Data Message . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Dilution Of Precision ( DOP ) . . . . . . . . . . . . . . . . GPS Error Sources . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Integrity / Navigation Requirements . . . . . . . . . . . GPS / DGPS Navigationaccuracy . . . . . . . . . . . . . . . . . GPS Operating Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Monitor Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GPS Predictive Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . HAM US/F-4 Sabrotzki Dec 2005 415 417 419 421 423 424 425 427 428 429 431 433 435 439 441 443 445 447 449 451 453 455 457 459 461 463 465 467 469 473 475 477 479 Page iv P66 B2 M13 4 NAV E TABLE OF FIGURES HAM US/F-4 Sabrotzki Dec 2005 Page v P66 B2 M13 4 NAV E TABLE OF FIGURES HAM US/F-4 Sabrotzki Dec 2005 Page vi

M13.02- Communication - Navigation- ATA 34

Related documents

Products

Support

M13.02- Communication - Navigation- ATA 34

Related documents

Add this document to collection(s)

Add this document to saved

Suggest us how to improve StudyLib