062 radio navigation (jaa atpl theory)

062 RADIO NAVIGATION

© G LONGHURST 1999 All Rights Reserved Worldwide

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EDITION 2.00.00 2001

This is the second edition of this manual, and incorporates all amendments to previous editions, in whatever form they were issued, prior to July 1999.

EDITION 2.00.00 © 1999,2000,2001 G LONGHURST

The information contained in this publication is for instructional use only. Every effort has been made to ensurethe validity and accuracy of the material contained herein, however no responsibility is accepted for errors ordiscrepancies. The texts are subject to frequent changes which are beyond our control.




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TABLE OF CONTENTS


Ground Direction Finding Stations

Automatic Direction Finding

VOR

The Radio Magnetic Indicator (RMI)

Distance Measuring Equipment

The Instrument Landing System

The Microwave Landing System

Basic Radar Principles

Ground Based Radars

Airborne Weather Radar



TABLE OF CONTENTS


Secondary Surveillance Radar

Area Navigation Systems

Doppler

Hyperbolic Navigation System Theory

Loran C

Satellite Navigation Systems



062 Radio Navigation



Loop Aerial Theory

Adcock Aerial

VDF Bearing Accuracy

Factors Affecting Range and Accuracy

VDF Approaches




Chapter 1 Page 1 © G LONGHURST 1999 All Rights Reserved Worldwide

1Ground Direction Finding Stations

1. Ground direction-finding (D/F) stations are normally located at airfields and enable air trafficto determine the bearing of an aircraft which is equipped with VHF radio (118 - 137 MHz), hencethe abbreviation VDF.

Loop Aerial Theory2. Figure 1-1 shows a vertical loop aerial, consisting of two vertical members, A and B,connected in the form of a loop by horizontal members. If a vertically polarised radio wave isincident upon the loop, it will induce voltages in the vertical members of the loop of value Va andVb. A current will therefore flow around the loop, the magnitude of the current being proportionalto the angle of incidence of the incoming radio wave.

FIGURE 1-1A Simple Loop Aerial





FIGURE 1-2Polar Diagram for a Vertical Loop Aerial

3. Figure 1-2 shows the horizontal polar diagram for a vertical loop aerial; it has two sharplydefined minima at θ== 0° and 270°, and two poorly defined maxima at θ ===0° and 180°.

4. If a loop aerial which is receiving a wave from a transmitter is rotated, the resultant voltage inthe loop will vary as θ=varies. When θ== 90° or 270° the resultant voltage is zero. When θ== 0° or180° the resultant voltage is a maximum. As the minima are the more sharply defined, these are usedfor direction finding. To take a manual loop bearing, the loop is rotated until a minimum signal, ornull, is found, when the transmitter must be on the line normal to the plane of the loop. However, itis not certain on which side of the loop the transmitter is sited. The process of resolving thisambiguity is known as sensing.





5. The loop aerial (in motorised form) is often used for airborne automatic direction findingpurposes, but is not best suited for ground installations. The wider the loop is, the better is its

direction finding ability (up to a maximum width of where is the wavelength of the received

signal). If the loop exceeds this width, its polar diagram becomes distorted. A loop aerial of this sizewould be impractical even for ground installation (150 metres wide for a frequency of 200 KHz), sosome form of fixed aerial is usually used for this purpose.

Adcock Aerial6. The simplest form of fixed aerial (Adcock aerial) consists of four uprights at the corners of asquare, with each diagonal pair joined by a screened cable which is often buried in the earth(Figure 1-3). At the centre of each screened cable is one of the stator coils of a goniometer - a devicewhich measures the direction of a magnetic field (see Chapter 2 paragraph 23 for a more detailedexplanation). A more effective form of the aerial is shown in Figure 1-4. This is known as the ‘H’type Adcock aerial. If the horizontal members in each ‘H’ are kept as close together as possible, thecurrents induced in each of the horizontal members will be identical and both will be in phase.Therefore there will be no current flow through the coil caused by induction in the horizontalmembers, so there will be no error due to down travelling or non-vertically polarised waves. Thecomplete aerial consists of two ‘H’ aerials crossing at right angles so that the coils in each circuitform the primary coils of a gonimeter.

λ10------ λ





FIGURE 1-3Adcock Aerial





FIGURE 1-4Adcock Aerial using Vertical Dipoles

7. As the pilot transmits, the sense directional aerial at the ground station receives the signal anddisplays the bearing to the air traffic controller, normally on a cathode ray tube (CRT) display suchas the one illustrated at Figure 1-5.





FIGURE 1-5Typical VDF Display

8. Appreciate that as this system operates in the VHF band it is therefore limited to line of sightconsiderations. The power of the signal transmitted from the aircraft will also limit the effectiverange at which a bearing is obtained.





9. The ground controller judges the accuracy of the bearing information given, on the basis ofthe length of the trace on the CRT screen, and classifies the bearing information accordingly.

10. Frequencies of ground stations which offer a VDF homer (QDM) service are listed in theCOM section of the AIP. There are additionally many automatic VDF stations whose function is toassist with ATC radar surveillance. In emergency they will provide DF assistance to aircraft (seelater), but the frequencies of these stations are not listed in the AIP.

11. The ground controller will give magnetic or true bearings, identified by the ‘Q’ code system asfollows;

(a) QDM, magnetic heading to steer to the station in zero wind conditions.

(b) QDR, magnetic bearing of the aircraft from the station.

(c) QTE, true bearing of the aircraft from the station.

(d) QUJ, true bearing of the station from the aircraft.

12. Of these QDM and QTE are most frequently used, although there is now a tendency to useQDR bearings rather than QDM bearings during the outbound leg of VDF approach procedures.

13. If a series of bearings is required by the pilot he should use the prefix QDL when requestingthe first bearing, for example:

G-LOST requests QDL QTE

14. The student will appreciate that if two or more (usually 3) D/F stations obtain bearings froma particular aircraft transmission it is possible to determine the approximate position of the aircraft -a process known as triangulation. Figure 1-6 shows the principle involved.





FIGURE 1-6Position Fixing using VDF Equipment

VDF Bearing Accuracy15. Bearings are classified according to their expected accuracy as follows:





16. In the UK bearing accuracy would not be expected to be better than Class B.

17. The pilot should be informed as to the class of bearing he is being passed by the air trafficcontroller, for example:

Your QTE is 278°, class bravo

Factors Affecting Range and AccuracyPropagation Error. At low altitude, where the signal is unevenly propagated over irregularterrain, the measured bearing may become distorted.

Multipath Signals. Reflections from buildings, etc, adjacent to the ground receiver may result inan inaccurate bearing being sensed.

Overhead Error. Accuracy is reduced when the aircraft position is close to, or directly overheadthe ground station.

Synchronous Transmissions. When two aircraft, within range of a ground receiver, transmit onthe same frequency at the same time, the resultant bearing information will be somewhere betweenthe two correct values.

Class A - ± 2°

Class B - ± 5°

Class C - ± 10°

Class D - more than ± 10°





Intervening Terrain. As mentioned earlier, ground D/F systems work at VHF and are thereforeline-of-sight systems. Any intervening high ground between the aircraft and the ground receiver mayresult in a shorter maximum D/F range depending on the relative heights involved.

Super-Refraction. Under certain meteorological conditions, radio waves in the VHF, UHF andSHF bands, which normally travel only in straight lines, may behave in a way which is at first sightsimilar to skywaves.

The meteorological conditions required for this alternative type of propagation (duct propagation)are a marked temperature inversion and a rapid decrease in humidity with height. Figure 1-7 showsducting which, in this case, is occurring between the surface and a low level inversion. The signal iseffectively trapped under the inversion and may travel hundreds of miles with little attenuation. Inthis way, when high pressure systems prevail, signals may be received from distant VHF transmitterswhich are far beyond the normal direct wave range.

FIGURE 1-7The Ducted Wave (Super Refraction)





The process of ‘ducting’ is also known as super refraction and as such may increase the range atwhich a ground D/F facility receives a given transmission.

Sub-Refraction. Sub-refraction is a condition of atmospheric refraction, created by gradients oftemperature and humidity, when radio waves are bent less than normal. This reduction in bendingwill result in a much smaller horizon distance than would be the case under normal propagationconditions. From a practical aspect a radio signal may appear to ‘fade away’ as the conditions occur.

VDF Approaches18. The following extract from the AIP is included for your guidance:

There are two types of VDF procedure, QDM and QGH. In the QDM procedure the pilot calls for aseries of QDM and uses them to follow the published approach pattern, making his own adjustmentto heading and height. In the QGH procedure the controller obtains bearings from the aircraftstransmissions, interprets this information and passes to the pilot headings and heights to fly designedto keep the aircraft in the published pattern. Normally, at civil aerodromes, only QDM procedure isavailable; however, in some cases, for specific operational reasons, there will be provision for QGHprocedure. Those aerodromes that have been approved to carry out both types of VDF procedurewill have this provision shown against the procedure. Pilots are reminded that it is their responsibilityto ensure with ATC that the correct procedure is being flown.






Non-directional Beacons

NDB Emission Characteristics

The ADF Receiver

The Automatic Function

The Bellini-Tosi System

The Control Panel

The Beat Frequency Oscillator

ADF Bearing Presentation

The Relative Bearing Indicator

Factors Affecting Range

Factors Affecting Accuracy

Procedures for Obtaining a Bearing





2Automatic Direction Finding

1. Automatic direction finding (ADF) or radio compass equipment is still widely used. Theairborne equipment (ADF) comprises a radio receiver designed to accept signals in the LF and MFbands, and to determine the great circle bearing of the transmitting station from the aircraft.

Non-directional Beacons2. Those transmitters which are specifically designed to give ideal signals for ADF use areknown as non-directional beacons (NDBs). These transmitters operate in the LF or MF bands, usingsurface wave propagation paths.

3. Many NDBs today are used to enable a pilot to locate either an airfield or the initial approachpoint of an instrument approach aid. Such beacons typically have a range of approximately 25 nmand are termed locator beacons. NDBs always transmit omnidirectionally.

4. In addition to using purpose built NDBs and locator beacons for navigation, the airborneADF equipment can utilise transmissions from any voice broadcast radio station that falls within therelevant frequency range.

5. The range of a surface wave is largely dependent upon the power of the transmitter. As aguide, a MF transmitter with a power of 10 kilowatts (kw) would have a range of about 500 nauticalmiles over water, over land somewhat lower ranges are likely. Relative power-to-range can beestimated from a rule of thumb which states that to double the range it is necessary to quadruple thepower.





6. As the wavelength of transmissions in the LF and MF bands is extremely large it becomesimpractical to utilise aerials which are the optimum length. In practice a convenient sized (butnevertheless quite large) aerial is used which is electronically ‘matched’ to the frequency in use (aprocess called ‘loading’). Typically two types of aerial are used with NDBs: the ‘T’ aerial(approximately 25m high and 50m long) for long range beacons, or tower aerials, approximately10m high and which are insulated from the ground. Both types of aerial are depicted in Figure 2-1 .





FIGURE 2-1Typical NDB Aerials





7. The NDB or locator beacon does not transmit a usable signal vertically upwards. Theinverted cone above the transmitter is known as the cone of silence or cone of confusion as indicatedin Figure 2-2. It should be noted that the period during which an aircraft will not receive usablesignals will increase as altitude is increased.

FIGURE 2-2Cone of Silence

NDB Emission Characteristics8. Two types of modulation characteristics are commonly used for non-directional beacons:





(a) NONA1A. The signal consists of two separate elements, the NON portion and theA1A portion. The NON part of the signal is, as we know, continuous carrier wave,and is ideally suited to enable the ADF to establish the direction from which the signalis arriving. The A1A part of the signal periodically replaces the NON transmissionand, being interrupted carrier wave, is used to carry the three letter morse identifierfor the NDB. NONA1A beacons are normally used with high power outputs for longrange NDBs.

(b) NONA2A. Similar to NONA1A stations but now the station identifier is carried bythe A2A signal (keyed single tone amplitude modulation). NONA2A beacons arenormally used for medium range NDBs.

9. Although most ADF receivers have a frequency selector range of 190 KHz to 1750 KHz, thefrequency bands which are internationally allocated to NDBs are 255 to 285 KHz and 315 to 405KHz.

The ADF Receiver10. The primary function of an ADF receiver is to determine the bearing of the incoming NDBsignal. Consider a loop aerial (Figure 2-3) which is connected to the aircraft's ADF receiver. Theloop is capable of rotating about a vertical axis. With the loop as shown at Figure 2-3, lying in theplane of the incoming signal, a current will be induced to flow through the loop by the NDB signalby virtue of the fact that one vertical element of the loop is further from the NDB than the othervertical element, and that consequently a phase difference exists between the two sides of the loopwith the result that there will be a voltage difference between the two vertical elements of the loop. Avoltage difference causes a current flow, which will be in the direction shown.





FIGURE 2-3ADF Loop Aerial Operation

11. Now consider the loop to be positioned as shown in Figure 2-4, lying at right angles to theincoming signal. The two vertical elements of the loop are equidistant from the NDB, and no currentwill flow through the loop since no phase difference exists between the two vertical elements.





FIGURE 2-4ADF Loop Aerial Operation (Cont’d)

12. Finally, consider the loop to be positioned as shown at Figure 2-5. The loop has rotatedthrough 180° from its original position and again a phase difference exists. Now a current will beinduced to flow through the loop, but as far as the receiver is concerned, in the opposite direction tothe first case considered.





FIGURE 2-5ADF Loop Aerial Operation (Cont’d)

13. The magnitude of current induced into such a coil would be very small indeed. This problemis overcome by using many thousands of coils wound on to a ferrite former, and by several stages ofamplification within the receiver.

14. Figure 2-6 shows the characteristic ‘figure of eight’ polar diagram which would be plotted ifthe signal strengths received by a loop aerial were plotted as a transmitter moved round the aerial ata fixed range through positions A,B, C and D. It will be seen that signal strength is maximum whenthe transmitter is at positions A and C, and zero when the transmitter is at positions B and D. The‘null’ produced in the latter cases is well-defined and can be used to determine the relative bearing ofthe transmitter from the aerial with reasonable accuracy.





FIGURE 2-6Polar Diagram of a Loop Aerial





15. Suppose the transmitter is at position D. If the loop aerial is rotated slightly in eitherdirection, signal strength will increase. By finding the null (zero signal strength) position, the bearingof the transmitter from the loop can be determined. However, there are two null positions at 180° toeach other and therefore an ambiguity exists, since the transmitter could be at either B or D.

16. To resolve this ambiguity a second, sensing aerial is added, which is designed so that thereceived signal produces an aerial current of the same strength as the maximum current in the loopaerial. Consequently, the radius of the polar diagram of this (single pole) aerial is equal to thediameter of each of the loop aerial circles, as shown at Figure 2-7.





FIGURE 2-7Polar Diagrams of Loop and Sense Aerials





17. The alternating current induced in one side of the loop will be out of phase with that in theother side (unless the loop is at right angles to the incoming radio wave), and this will result in apotential difference between the two sides of the loop, and a resulting current flow. This isrepresented by +ve and -ve signs in the loop polar diagram (Figure 2-7). The sensing aerial inducedcurrent is of constant phase and will consequently be in phase with the loop aerial if the transmitteris in position A (for example), and 180° out of phase if the transmitter is at position C.

18. The combined polar diagrams of the sense and loop aerials will depend upon algebraicaddition of their signs. This produces the cardioid (heart shaped) polar diagram as shown atFigure 2-8. It will be seen that the combined polar diagram is used to resolve the 180° ambiguity.





FIGURE 2-8ADF Combined Polar Diagram





19. Referring to Figure 2-8, suppose the transmitter is at position D. The loop aerial senses anull, which means the transmitter must either be at D or B. If the loop is rotated anti-clockwise thecombined polar diagram will rotate with it and signal strength will increase. If the loop is rotatedclockwise, signal strength will decrease. If the transmitter were in position B, the reverse would betrue. Thus, if anti-clockwise rotation produces a stronger signal, the bearing is correct, however if itproduces a weaker one the bearing is a reciprocal.

20. Hence, the bearing of the transmitter can be found by manually rotating a loop/sense aerialcombination and listening to the strength of the received signal.

The Automatic Function21. Figure 2-9 shows a schematic diagram of an automatic direction finding system (in fact aBellini-Tosi system, discussed shortly). As already described, if the transmitter is to the left of theloop ‘null line’ (the line joining the two null bearings, for example the line joining points B and D atFigure 2-6 and Figure 2-8), the sense aerial current is in phase with the loop aerial current. If thetransmitter is to the right, the two currents will be out of phase.

22. A two phase motor is used to drive the loop and the bearing indicator. If the two aerialcurrents are in phase the motor will rotate the aerial clockwise until the loop aerial senses a null. Ifthe two signals are in antiphase the motor will rotate the loop anti-clockwise until the null of theloop polar diagram is sensed.





FIGURE 2-9Automatic Direction Finding System. Fixed Loop Installation.





The Bellini-Tosi System23. Modern ADF systems work on exactly the same principle as that described above. In high-speed aircraft it is not desirable to mount a bulky loop aerial housing outside the fuselage. Toovercome this problem two coils, which are at right angles to each other, are wound on ferrite cores,embedded in a flat block of insulating material and fitted flush with the aircraft skin. The two loopsare connected to the stator coils of a synchro called a goniometer. The field produced by theincoming signal is now effectively reproduced within the goniometer. A search coil lying within thefield produced by the goniometer stator coils now rotates in search of the null position; the ADFpointer is controlled by the movement of this search coil. The system is shown diagrammatically atFigure 2-9.

The Control Panel24. Figure 2-10 shows the face of a modern ADF receiver. This receiver is relatively devoid ofcontrols and, compared with the older sets, is very easy to use. The frequency selection controlsenable the operator to select the frequency required by placing the appropriate numbers in thefrequency window. Note that, like most modern equipment, this unit has 0.5 KHz frequencydivisions.

25. The function switch has four positions. With ADF selected and a suitable signal present theequipment automatically gives bearing information.

26. By selecting the ANT (antenna) position the loop aerial is taken out of circuit and theaudibility of A2A NDB identifiers should improve. Obviously with the ANT position selected thebearing information must be disregarded, but it is the best position for checking signal strength andidentification.





FIGURE 2-10Typical ADF Control Panel.

27. By selecting the BFO (beat frequency oscillator) position, an oscillator within the ADF isbrought into the circuit, and creates an audible output from non amplitude modulated (NON andA1A) inputs. This enables the pilot to check the NON portion of the incoming signal for fading(night effect) or high noise levels (thunderstorm effect or precipitation static), and of course to checkthe ident of an NDB using A1A modulation. Again, since the loop aerial is isolated when the ADFreceiver is in the BFO mode, the bearing must be ignored whilst the function switch is in the BFOposition. On some equipments the BFO is a separate switch, independent of the function control.





28. The Bandpass Selector Switch (Broad/Sharp Switch) allows the operator to narrow the bandof frequencies fed to the receiver if it is necessary to exclude unwanted background noise. It shouldbe noted that Broad (or wide bandwidth) should be selected when listening to voice or music.

29. Depressing the test button causes the ADF needle to swing a preset amount (typically at least90°). When the button is released the needle should swing back to the original reading. Olderequipments incorporate a loop control facility with which a similar check can be made by rotating

the loop first clockwise and then anti-clockwise through at least 900.

The Beat Frequency Oscillator30. An amplitude modulated signal is demodulated in a conventional receiver without anydifficulty since the amplitude of the carrier wave is varying in sympathy with the intelligencewaveform. With a NON or A1A signal the amplitude of the carrier wave remains constant andtherefore it is impossible to achieve an audible output from a receiver using conventionaldemodulation techniques.

31. Figure 2-11 shows how the receiver is modified when the BFO function is selected. The BFOis made to generate an alternating current, the frequency of which differs from the incoming carrierwave frequency by, typically, 2 KHz. The incoming signal and the BFO-generated signal are both fedto the heterodyne unit (frequency mixing unit) which gives four output frequencies. The output ofthe heterodyne unit comprises the two input frequencies, the sum of the two input frequencies, andthe difference frequency. It is only the difference frequency (2 KHz) which is audible, and this is fedto the loudspeaker, producing the audio tone.





FIGURE 2-11Receiver with BFO Facility





ADF Bearing Presentation32. ADF bearing information is presented to the pilot either as a relative bearing on a RelativeBearing Indicator (RBI), or as a magnetic bearing on a Radio Magnetic Indicator (RMI).

The Relative Bearing Indicator33. The RBI is used solely for ADF bearings, and is illustrated at Figure 2-12.

FIGURE 2-12Relative Bearing Indicator

34. The information shown on the RBI at Figure 2-12 is simply the angle subtended between theaircraft nose (zero degrees relative) and the path of the incoming NDB signal. By convention thepointed end of the needle always points towards the NDB.





35. In order to convert a relative bearing at the aircraft into a true bearing to plot from the NDBthe following procedure is adopted:

(a) Convert the aircraft compass heading to a true heading by applying deviation, andvariation at the aircraft position.

(b) Add the true heading to the relative bearing.

(c) Add or subtract 180°.

36. The result of the above calculation will be the true great circle bearing of the aircraft from theNDB. No consideration has been made here of convergency or conversion angle.

37. Figure 2-13 and Figure 2-14 illustrate the procedure outlined above for plotting relative ADF/NDB bearing (ignoring convergency and conversion angle).





FIGURE 2-13Calculation of Bearings using RBI

38. A far more convenient way of presenting ADF bearing information is on the Radio MagneticIndicator (RMI). If an aircraft is fitted with RMIs it is normal to present both ADF and VORbearings on the same instrument. You will find an in-depth consideration of the RMI in chapter 4.





FIGURE 2-14Calculation of Bearings using RBI (large numbers)

Factors Affecting Range39. The points discussed below are the major factors which will determine the maximum range atwhich satisfactory bearings may be obtained:

(a) NDB transmitter power. The greater the power output of the NDB the greater therange.





(b) NDB frequency. As discussed in the previous section, the lower the frequency thesmaller the rate of surface wave attenuation and consequently the greater the range.

(c) The surface. Greater range is achieved over the sea than over the land.

(d) Type of emission. NONA1A emissions, having a very narrow bandwidth, give longerranges for a given power output than NONA2A or A2A emissions, with their broaderbandwidth.

(e) Precipitation static. Electrical discharges which occur when precipitation strikes theairframe will increase the ambient radio noise level, and this may be sufficient toobscure the incoming NDB signal, thereby limiting the range at which an NDB isusable. Static wick dischargers mounted on the trailing edges help to discharge theaircraft's static electricity to atmosphere, thereby minimising the effect.

40. As the range of the aircraft from the NDB in use increases, so the signal becomes weaker, andtherefore the signal to noise ratio decreases. In the United Kingdom, the minimum signal to noiseratio which is considered acceptable is 3:1, and this should produce a bearing accuracy of within ± 5°by day only, within the promulgated range.

Factors Affecting Accuracy41. ADF systems suffer from a number of errors, which are discussed below.





Dip Error42. The operation of the loop aerial as described on page 2-3 is dependent upon voltagedifferences that are created only in the vertical elements of a loop aerial. However, when bank isapplied to an aircraft the horizontal arms of the loop aerial will tilt such that they have a verticalcomponent to their orientation; thus they will also have a current flow induced in them and thisadditional current will result in a small error in indicated bearing. This error is known as dip error.

Mountain Effect43. Hills and mountains reflect and re-radiate the LF or MF signals of an NDB. Consequently,the ADF in a low-flying aircraft may receive both the great circle signal and a re-radiated signal at thesame time. This will result in an erroneous bearing, as shown at Figure 2-15.

FIGURE 2-15Mountain Effect





Quadrantal Error44. The airframe itself tends to reflect, refract and re-radiate the incoming signal. The result isthat the loop aerial will receive a signal directly from the NDB together with a much weaker signalwhich has been distorted by the aircraft fuselage. The net effect is that the incoming signal appearsto bend towards the fuselage as illustrated at Figure 2-16. Notice that it is signals arriving on relativequadrantal bearings which are most affected. Signals arriving on relative cardinal bearings are notnormally affected to any significant degree. It is possible to calibrate quadrantal errors out of thesystem when the ADF is installed into the aircraft.

FIGURE 2-16Quadrantal Error





Coastal Refraction45. Because radio waves travel marginally faster over the sea than over the land, any radio wavecrossing the coastline at other than 90° will be refracted. Figure 2-17 shows the waves being bentaway from the normal when crossing from land to sea. Notice that the further the signal is from thenormal, the greater the amount of refraction.





FIGURE 2-17Coastal Refraction





46. To minimise coastal refraction, bearings should be taken when the aircraft is positioned suchthat the signal from the NDB is crossing the coast at an angle fairly close to 90°. If this is notpossible an NDB should be used which is as close as possible to the coast. Figure 2-18 illustrates whythis is so.





FIGURE 2-18Reduction of Coastal Refraction.





Night Effect47. Within the LF and MF bands skywaves are not normally present by day, since the ionosphereis intensely ionised and totally attenuates all LF and MF radio waves entering the layers. By night theionosphere is partially de-ionised and now NDB signals may survive to be refracted back to thesurface of the Earth.

48. What is now happening is that both surface wave and skywave signals from the NDB in usemay arrive at the aircraft together. It is likely that the two waves will be out of phase. Additionally,if the ionosphere does not lie parallel with the Earth's surface, the two signals will arrive at theaircraft along different great circle paths. The net result is that the ADF bearing will be in error.

49. In fact night effect is most pronounced during the twilight periods, since at these times theionosphere is changing both its intensity and its height above the surface.

50. When night effect is affecting the incoming signal the needle will tend to wander, and theidentification signal to fade, as the two incoming signals (the surface wave and the skywave) drift inand out of phase with each other.

51. Appreciate that night effect is assumed to occur because of the inter-action of the surfacewave from the NDB in use and the skywave, also from the NDB in use. Obviously, skywaves fromNDBs or other stations operating on the same frequency will affect the accuracy of the bearing (seestation interference), however this distant station skywave interference is not considered to be nighteffect, even though it will be most pronounced at night.





Station Interference52. In order to ensure that there is little or no interference between NDBs operating on the sameor similar frequencies, both beacon location and frequency allocation are carefully planned. Surfacewave coverage of NDBs on the same frequency should not therefore overlap. If it is not possible tototally prevent this surface wave overlap situation, the NDBs concerned are given promulgatedranges. This range, which is published in the AIP, denotes the maximum range at which the NDBsignal should be considered as being free from harmful distant station interference BY DAY.Figure 2-19 illustrates the significance of promulgated range.





FIGURE 2-19Station Interference/ Promulgated Range





Static Interference53. Static interference is one of the largest sources of error in the operation of NDB/ADF systems.All kinds of precipitation (including falling snow) and thunderstorms can cause static interference ofvarying intensity. Precipitation static, described in the Radio Theory part of the ATPL syllabus,reduces the effective range and accuracy of bearing information. Thunderstorm activity can give riseto bearing errors of considerable magnitude and even to false ‘overhead’ indications. The electricalemissions during a thunderstorm may well result in the ADF indicating the direction of the stormrather than the NDB

54. By night the skywaves from distant stations may well reach the aircraft, even though it isoperating well within the promulgated range of the beacon in use. It is for this reason thatpromulgated ranges are not valid by night. Figure 2-20 illustrates this very important fact.





FIGURE 2-20Problems Associated with Promulgated Range

Lack of Failure Warning System55. The majority of ADF instruments do not incorporate a failure warning indication.Consequently, failure of any part of the airborne receiving or ground transmitting apparatus mayproduce false bearing indications which are not readily detectable.





56. In particular, failure of the NDB could adversely affect both systems of a dual ADFinstallation in the aircraft, leading to the false assumption that - because both are indicating the same- both are correct.

57. In order to reduce the risk of a false indication being followed, the correct method ofassessing system performance is to continuously monitor the NDB audio identification signal and theRMI/RBI pointer behaviour. This is particularly applicable when making an approach toward theNDB when, in the event of failure, the pointer could give a reverse indication. Since this should onlyoccur once the beacon has been passed, reversal or marked change at any other time may be taken asan indication of probable system failure.

58. Loss of the NDB identification signal may be taken as an indication of NDB failure, sinceunder these circumstances the identification signal is suppressed, or replaced by a continuous tone.

Procedures for Obtaining a Bearing59. That which follows is a full and correct procedure for obtaining a bearing from an NDBclassified as NONA2A.

(a) Before flight, select the NDBs required for flight navigation and check the AIP forpromulgated range, modulation characteristics and scheduled servicing periods, andthe current Notams for frequency or location changes, unserviceability or non-scheduled servicing periods.

(b) Before take-off, check serviceability of ADF receiver using two NDBs of knownbearing from the airfield.

(c) In flight, select the required frequency.





(d) With the function switch at ADF or ANT, check the Morse ident, and ensure thatthere is no A2A station break-through.

(e) Select function switch to BFO. Check for steady DF tone, with no fading (nighteffect), no high noise level (thunderstorm effect), and no A1A station break-through.

(f) Select function switch to ADF and ensure that the needle points steadily in what youconsider to be approximately the correct direction.

(g) Deflect the needle using the press to test button, release, and ensure that the needlereturns to its original position.

(h) Note bearing, heading (for RBI readings only), and time.





Self Assessed Exercise No. 2

QUESTIONSQUESTION 1.

Under certain meteorological conditions radio waves in the VHF, UHF and SHF bands may bereceived at ranges far beyond the normal direct wave range. This phenomenon is called _________ .

QUESTION 2.

A class C VDF bearing would have an expected accuracy of _________ .

QUESTION 3.

The range which might be expected from an NDB which is transmitting 10 Kilowatts of power overthe sea in average conditions is:

QUESTION 4.

The principal propagation path of an NDB is:

QUESTION 5.

What types of aerial are used with NDBs?

QUESTION 6.

The Loop aerial in an ADF is used for:





QUESTION 7.

When the induced signals from the loop and sense antennae are combined in an ADF receiver, theresultant polar diagram is:

QUESTION 8.

What frequency selection range does ADF equipment cover?

QUESTION 9.

What is the purpose of the Bandpass Selector Switch (Broad/Sharp switch) on an ADF control panel?

QUESTION 10.

In an ADF receiver, night effect is most pronounced.

QUESTION 11.

When ANT is selected on the ADF function switch which aerial(s) are being used?

QUESTION 12.

In an ADF receiver, thunderstorm effect is caused by:

QUESTION 13.

Does the Beat Frequency Oscillator in ADF equipment generate an alternating current at an audio orradio frequency?





QUESTION 14.

An NDB bearing is taken from an RBI and reads 057° relative. The heading of the aircraft is 359°(C),deviation is 3°E and magnetic variation is 7°W. What is the calculated true bearing to plot from theNDB to the aircraft position?

QUESTION 15.

What is the quoted bearing accuracy of an NDB?

QUESTION 16.

How is Quadrantal Error corrected in an ADF system?

QUESTION 17.

Other than taking bearings that cross the coast at right-angles, how may the effects of CoastalRefraction be reduced when using ADF equipment?

QUESTION 18.

What is the validity period of promulgated ranges?

QUESTION 19.

Does Coastal Refraction have the effect of bending a radio wave towards or away from the normalto the coastline, when the signal is travelling from land to sea?





QUESTION 20.

The RBI bearing of an NDB is 343° relative. The heading of the aircraft is 350°(M), deviation is 4°Wand magnetic variation is 5°E. What is the calculated true bearing to plot from the NDB to theaircraft position?

ANSWERS:ANSWER 1.

Super refraction (or ducting)

ANSWER 2.

±10°

ANSWER 3.

500nm

ANSWER 4.

Surface Wave

ANSWER 5.

T aerial or Tower aerial.

ANSWER 6.

Direction finding using the null positions of the polar diagram.





ANSWER 7.

Cardioid

See FIGURE 209 in the Reference Book

ANSWER 8.

190 Khz to 1750 Khz

ANSWER 9.

It reduces the bandwidth of the receiver to exclude unwanted noise.

ANSWER 10.

At dawn and dusk

ANSWER 11.

With ANT selected only the Sense Aerial will be in use.

ANSWER 12.

The ADF receiver being attracted to naturally occurring electromagnetic radiation.

ANSWER 13.

The BFO is a RADIO FREQUENCY oscillator





ANSWER 14.

Hdg = 359°(C) Dev = 003°E Mag Hdg = 002°(M) Varn = 007°W True Hdg = 355°(T)

Brg = 057°(R) True Hdg = 355° = 412° - 360 = 052°(T)Brg from NDB = 180 + 052 = 232°(T)

ANSWER 15.

±5° by day only (within the promulgated range).

ANSWER 16.

A Quadrantal Error corrector is installed with the ADF system and the error is calibrated out.

ANSWER 17.

By using NDBs that are close to the coastline





ANSWER 18.

Promulgated Ranges are only valid by day due to the possibility of skywave interference at night.

ANSWER 19.

The radio wave will be bent away from the normal to the coastline.

ANSWER 20.

Mag Hdg = 350° Varn = 5°E True Hdg = 355°

NDB Brg = 350° True Hdg = 355° = 698° - 360° Brg to NDB = 338°(T)

Brg to NDB = 338°(T) - 180°Brg from NDB = 158°(T)





VOR

Principle of Operation

VOR Frequency Range and Chart Symbology

Operational Range of VOR

VOR Errors

Types of VOR Transmitters

VOR Bearing Presentation



VOR


3VOR

1. VHF Omnidirectional Radio Range (VOR) is a system which gives accurate bearings withreference to ground-based stations.

Principle of Operation2. VOR stations transmit a carrier wave which is modulated in a manner previously described asA9W in the Radio Theory part of the syllabus. This is to say that the single carrier wave is bothfrequency and amplitude modulated at the same time.

3. By frequency modulating the carrier wave with a simple 30 Hz wave form, the referencesignal is achieved. This signal is so named since all airborne receivers at a given range from thestation will receive a reference signal which is at the same phase, regardless of the aircraft bearingfrom the station.

4. The VOR station transmits in all directions (omnidirectionally), however the signal strengthvaries depending on the bearing from the station at a given point in time. The polar diagram for theVOR transmitter, which is known as a limacon, is illustrated at Figure 3-1.



VOR


FIGURE 3-1Limacon Polar Diagram

5. The limacon itself is rotated at the rate of 30 revolutions per second and this has the effect ofamplitude modulating the carrier wave arriving at an airborne receiver, see Figure 3-2.

6. At Figure 3-2(a), the aircraft to the east of the beacon is receiving a signal of minimumstrength, since the shortest radius of the limacon is facing the aircraft.

7. At Figure 3-2(b), the amplitude of the signal arriving at the aircraft is of mean value, since thelimacon has rotated through 90°.



VOR


8. At Figure 3-2(c), the signal is at maximum amplitude, the limacon having rotated through afurther 90°.

9. At Figure 3-2(d) the signal is back to a mean value, and at Figure 3-2(e) the signal is atminimum strength, since the limacon has now rotated through one complete revolution.



VOR


FIGURE 3-2Amplitude Modulation Produced by Rotating the Limacon



VOR


10. The phase of the amplitude modulated signal will depend upon the bearing of the airbornereceiver from the ground station. The amplitude modulation is therefore known as the variphasesignal.

11. Each airborne receiver is now receiving both a reference and a variphase signal and therotation of the limacon is so arranged that the two signals are in phase to any observer on a magneticbearing of 360° from the VOR beacon. The phase difference between the reference and variphasesignals will now relate directly to the magnetic bearing of the receiver from the beacon, seeFigure 3-3.



VOR


FIGURE 3-3Signal Production in a VOR



VOR


VOR Frequency Range and Chart Symbology12. VOR operates in the VHF band between 108.0 MHz and 117.95 MHz. The frequenciesallocated to VOR within this band are:

(a) 108.0; 108.05; 108.2; 108.25; 108.4; 108.45 and so on up to 111.8 and 111.85 MHzgiving 40 channels;

(b) 112.0; 112.05; 112.1; 112.15; 112.2 and so on at 0.05 MHz (50 KHz) spacing to117.9 and finally 117.95 MHz. MHz, giving a further 120 channels.

13. The reason for the gaps (for example 108.1 and 108.15 MHz) between 108 and 112 MHz isthat the 40 frequencies allocated to the ILS Localiser also lie between 108 and 112 MHz.

14. All VORs transmit a three-letter morse identification code (1020 Hz amplitude modulation),which is repeated six times a minute, unless the VOR is paired with a DME, but more of that later.When a VOR station is transmitting for test or calibration purposes, the normal three letteridentification is replaced with the three letters ‘TST’ (see later). When this occurs the VOR must notbe used for navigation.

15. Chart symbology for VORs varies from chart to chart. Figure 3-4 shows the most commonsymbols used.



VOR


FIGURE 3-4VOR Chart Symbols



VOR


16. When plotting a VOR radial on a navigation chart the reciprocal of the bearing taken froman RMI (see chapter 4) should be noted, and converted to a True bearing by applying MagneticVariation at the beacon location. This True bearing should then be plotted, with respect to TrueNorth, from the VOR position.

Operational Range of VOR17. VOR operates in the VHF band and all transmissions are therefore limited to line of sight.

18. The power output of the VOR station will also affect the operational range. En-route VORsnormally use 200 watt transmitters giving a range in excess of 200 nm. Low power VORs are oftenused as airfield beacons, and these stations normally transmit only 50 watts, giving a range of 100nm or so.

VOR Errors19. The VOR station produces a complex A9W carrier wave conveying magnetic bearinginformation. If either the reference or the variphase signal shifts, the signals are no longer in phasewith each other at magnetic north. Consequently the bearing information is now erroneous. Thiserror is due to ground equipment malfunction. Such an error is not allowed to persist, since amonitor station always forms part of the ground installation.

20. If the monitor senses an error in excess of one degree at the transmitter, a standby transmitteris automatically brought to a state of readiness. This takes several minutes, but as soon as thestandby transmitter is ready a changeover is automatically achieved. During the period when thestandby transmitter is preparing for the changeover, the main transmitter continues to transmiterroneous bearing information. To warn the pilot that the bearing information is unreliable, themorse identifier is suppressed. During this changeover period it is unlikely that an alarm flag will



VOR


appear on the VOR indicator at the aircraft. In this case an omission of the morse identifier is theonly indication to the pilot of system malfunction. The monitor will also initiate a change to thestandby VOR transmitter in the event that a drop in the strength of the radiated signal from themain transmitter of 15% or more is sensed. Again the identifier will be suppressed during thechangeover period. Finally, the morse identifier will be removed from the carrier wave in the eventthat the monitor fails.

21. Having established that the ground equipment error will not exceed one degree, it is nownecessary to consider other errors which may degrade the accuracy of the bearing displayed at theaircraft.

Site Errors22. Figure 3-5 shows that transmitted energy from the beacon may reach the aircraft viareflecting surfaces such as hills or nearby buildings, as well as along the direct path. The compositesignal reaching the aircraft in this case will have an electrical phase that differs from the direct wavephase and this will result in significant bearing errors. To minimise site errors the transmitter aerial issited, whenever possible, on flat terrain in an area remote from buildings or hilly ground.



VOR


FIGURE 3-5Site Errors

Airborne Equipment Errors23. It is impossible to measure the phase difference between the reference and variphase signalswith total accuracy. Each one degree of error in measuring the phase difference results in a one-degree error in bearing information. This phase comparison error at the aircraft should not exceedplus or minus three degrees (±3°).

Propagation Errors24. Propagation errors, otherwise known as scalloping, may occur at extreme range and at lowaltitude where the signal is unevenly propagated over irregular terrain.



VOR


25. If a VOR beacon is known to be subject to either site error or propagation error, thisinformation will appear in the remarks column for that particular VOR in the COM section of theAIP. From a practical point of view, beam bends and scalloping of VOR bearings/radials oftenmanifest themselves when flying along an airway centreline between two VORs. The aircraft may betracking the centreline quite happily using the bearing indications from the VOR behind the aircrafthowever, when the pilots retune the VOR to the beacon ahead of the aircraft, the new indicationscould show a relatively large deviation from track.

Station Interference26. If an aircraft is operating at altitude the situation may arise whereby the VOR set is receivingsignals from two stations operating on the same frequency. One of these stations will presumably bethe station selected by the pilot, the other station will therefore be interfering with the desired signal.The consequence of this station interference will be an erroneous bearing indication.

27. With only a limited number of spot frequencies allocated for use by a great number of VORstations, careful planning is essential if station interference errors are to be avoided. This is primarilyachieved by wide geographical spacing of VOR beacons using the same frequency.

28. When it proves impossible to prevent a degree of signal overlap, a Designated OperationalCoverage (DOC) is published in the AIP for the VOR stations concerned. The DOC defines thevolume of airspace within which harmful interference from distant stations is avoided. Unlike thepromulgated range of NDBs discussed previously, DOC is valid both by day and by night, sinceskywaves do not occur in the VHF band at any time.

29. In some cases interference may be limited to one sector, or may be more significant in onesector than in others. This is illustrated in the extract of the COM section of the AIP shown atFigure 3-6.



VOR


FIGURE 3-6Example of Designated Operational Coverage

30. The table at Figure 3-7 summarises the errors likely to affect a VOR bearing.

FIGURE 3-7Summary of Errors in a VOR

Types of VOR Transmitters31. Described below are the various types of VOR station in common use.

Station Service Callsign/Ident Frequency Remarks

Burnham VOR BUR 117.10 MHz DOC 60 nm/ 25,000 ft. Ignore any DME indications, no associated DME.

Dover VOR/DME DVR 117.70 MHz DOC 80 nm/50,000 ft (200 nm/50,000 ft in sector 025° - 039°(M). DME on channel 124X

Transmitter error - monitored to remain within

Airborne equipment error - typically a maximum of

Station interference - within the DOC should not exceed

Site error - insignificant with careful siting (or Doppler VOR)

Propagation error - insignificant if the VOR is used at sensible range and altitude

± 1°

± 3°

± 1°

-

-

Total ±5°



VOR


Standard VOR32. Widely used to define airways centrelines, transmitter power normally 200 watts.

Terminal VOR (TVOR)33. Low-powered beacons used as airfield location aids. Widely used in some parts of the world,in conjunction with DME, as a procedural approach aid.

Broadcast VOR (BVOR)34. Broadcast VOR beacons provide bearing information in the usual manner. Additionally, avoice modulation is superimposed on the carrier wave. The audio information normally provides anAerodrome Terminal Information Service (ATIS) giving present weather, runway in use andserviceability state of the airfield and associated navigation aids.



VOR


Test VOR (VOT)35. A VOT is a VOR, which operates on one of the normal VOR frequencies, and provides a testsignal for the pre-flight checking of VOR airborne equipment. The VOT radiates an RF carrier onwhich are superimposed two separate 30 Hz modulated waveforms which are identical in format tothe reference and variphase signals of a conventional VOR. Both modulating waveforms are ‘phaselocked’ together with 180° phase difference such that, no matter where an aircraft is positioned inrelation to a VOT, the indications on the flight deck will always show the aircraft to be on the 180°radial from the ground facility. Identification of the VOT beacon is done in the normal mannerhowever, the appropriate authority will ensure that the ident sequence is unmistakenly distinctive asto the test function (in certain areas, where VOT coverage is confined to a single aerodrome, theidentification may consist of a series of dots). The accuracy of the VOT test bearing should be within± 1°.

Doppler VOR (DVOR)36. VOR transmitter aerials should be sited on flat terrain to minimise site errors. If such a site isnot available, a complex aerial system may be employed to transmit the VOR signal. This type ofstation is known as a Doppler VOR (DVOR) beacon and produces a signal which is reasonably freeof site errors even when the transmitter is sited in hilly terrain.

37. The way in which the bearing signal is produced is quite different from conventional VOR,the received signals are indistinguishable from each other and the airborne receiver will operate oneither with equal facility. In doppler VOR the reference signal is amplitude modulated at 30 Hz,whilst the bearing signal is frequency modulated at 30 Hz. Because this is the reverse of conventionalVOR, the bearing (or variable) modulation is made to lead the reference signal by a phase angleequal to the aircraft's magnetic bearing from the VOR ground station.



VOR


38. The Doppler VOR transmitter comprises a circle of about 50 antennae surrounding a singleomni-directional antenna. The latter transmits the AM reference signal, whilst the circle of antennaeare sequentially energised in an anti-clockwise direction at 30 revolutions per second (30 Hz). Fromany given direction, it will appear as though the transmitter is advancing and retreating at 30 Hz - inother words there will be a Doppler shift. The phase relationship between the doppler shift and thesteady reference signal is arranged to be zero when received on a bearing of 0°(M) from thetransmitter. Since both signals have the same modulating frequency (30 Hz), at 180°(M) from theVOR the phase difference will be 180°, at 270°(M) it will be 270° and so on.

VOR Bearing Presentation39. VOR bearing information is presented to the pilot either on an Omni Bearing Indicator(OBI), or on a Radio Magnetic Indicator (RMI).

The Omni Bearing Indicator40. The basic Omni Bearing Indicator (OBI) is illustrated at Figure 3-8. The instrument is oftendesigned to serve a dual function as both VOR bearing indicator and ILS meter, depending upon thefacility selected. For clarity, the ILS glidepath needle is omitted from all illustrations in this chapter.



VOR


FIGURE 3-8Omni Bearing Indicator

41. Check the illustration at Figure 3-8 and note the following points:

(a) The instrument face illustrated is known as a five-dot display. The ring at the centrerepresents the first dot.

(b) In the VOR mode the track deviation needle will deviate from the centre of theinstrument by one dot for every two degrees that the aircraft is displaced from themagnetic track selected in the window. Full-scale deflection of the needle thereforerepresents a track error of ten degrees or more.



VOR


(c) The needle represents the required magnetic track. The centre of the instrumentrepresents the aircraft.

(d) The required magnetic track is selected in the window using the omni bearing selector(OBS) knob.

(e) A prominent alarm flag will appear whenever:

(i) The airborne receiver fails, or power supply is lost.

(ii) The aircraft receives no acceptable VOR signal, due to range, height, orbecause the aircraft is directly overhead or abeam the station.

(iii) The VOR ground station fails.

(f) When the alarm flag is not visible it will be replaced by either a TO or a FROMindication depending on the aircraft position in relation to the VOR station and to themagnetic track selected in the window.

42. The VOR station does not transmit a usable signal vertically upwards. The inverted coneabove the transmitter is known as the cone of silence or the cone of confusion. The ICAOrequirement is that the cone subtends an angle of not more than 50° from the vertical, as shown atFigure 3-9. The period during which an aircraft will not receive usable signals will increase asaltitude is increased. During this period the alarm flag will be visible, indicating station passage.

43. The maximum radius or diameter of the cone of silence can easily be calculated for a givenaltitude using simple trigonometry or scale drawing.



VOR


FIGURE 3-9Cone of Silence

44. The alarm flag will also be visible for a period of time if the aircraft flies abeam the beacon(with respect to the magnetic track selected). The two areas of ambiguity (the two abeam sectors)extend through 20° arcs equally displaced about the perpendiculars to the magnetic track selected,see Figure 3-10.



VOR


FIGURE 3-10Areas of Ambiguity



VOR


45. Remember that the OBI does not think for itself, and it has no knowledge of aircraft heading.The OBI will present positional information relative ONLY to the MAGNETIC TRACK selected inthe window.

46. Please re-read the last paragraph and make a mental note not to fall into the trap either in theexamination or in the air. Figure 3-11 and Figure 3-12 illustrate the problem.

FIGURE 3-11Correct Track Selection on an OBS



VOR


FIGURE 3-12Incorrect Track Selection on an OBS

47. At Figure 3-11 the aircraft is attempting to track 300°(M) outbound from the VOR stationand the required magnetic track of 300° has been correctly dialled in the window. The aircraft ispresently located on the 306° radial, the OBI shows three dots fly left and the FROM flag is visible.All is as it should be.

48. At Figure 3-12 the aircraft is attempting to track 120° inbound to the VOR station (to trackinbound on the 300° radial). The reciprocal of the required magnetic track has been dialled in thewindow. The aircraft is again located on the 306° radial. The needle shows three dots fly left whenit should show three dots fly right. The FROM flag is visible and it should be the TO flag. All isNOT as it should be.



VOR


49. In order to obtain a bearing to plot from a VOR station to the aircraft using an OBI simplyrotate the OBS until the needle is vertical with the FROM flag showing. The reading in the trackwindow can now be plotted using magnetic north at the station. Alternatively variation at thestation can be applied and the resultant bearing plotted from the meridian passing through thestation.

50. It is not necessary to apply convergency when plotting VOR bearing on a Lambert or polarstereographic chart. It is however necessary to apply a correction for conversion angle when plottingVOR bearings on a Mercator chart, if the change of longitude between aircraft and station issignificant.

51. Using the OBI for tracking is a straightforward exercise. Select the required magnetic track inthe window and keep the needle in the middle, making necessary allowance for drift. Appreciate thatthe basic OBI described shows angular displacement from track, and consequently the correspondinglinear displacement expressed in terms of distance will depend upon aircraft range from the station.For example (using the 1 in 60 rule), two dots fly left at a range of 60 nm represents a track error of4 nm. The same two dots fly left at a range of 6 nm represents a track error of less than half anautical mile.






In-flight ADF and VOR Procedures using an RMI





4The Radio Magnetic Indicator (RMI)

1. The RMI is used both for ADF and VOR bearing information, and is illustrated at Figure 4-1.





FIGURE 4-1RMI Presentation





2. The ADF needle on the RMI is still effectively showing the relative bearing of the NDB (theNDB bears 090° relative to the aircraft nose, and the sharp end of the needle is 90° removed from thetop of the RMI). The difference is that the compass rose on the RMI is slaved to the aircraft gyro-compass and presently indicates the aircraft heading of 060°(M) at the top of the instrument. TheRMI has therefore mechanically added the magnetic heading to the relative bearing and the needleshows the magnetic bearing of the NDB from the aircraft.

3. The VOR bearing is displayed on the RMI with the sharp end of the needle pointing to theVOR station.

4. It is probable that the compass to which the RMI is slaved will suffer small amounts ofdeviation. The amount of any deviation should be very small (certainly less than one degree for asophisticated gyro-slaved system) and is normally ignored. Should the amount of deviation becomesignificant it is necessary to correct ADF bearings on the RMI; VOR bearings are not affectedbecause the equipment first subtracts aircraft heading to produce a relative bearing which controlsthe pointer, but on the display aircraft heading is re-applied to give the original QDM as the output.





FIGURE 4-2The Effects of Magnetic Deviation





5. Figure 4-2 shows the same situation as given at Figure 4-1, however now 10° east deviation(+10°) has been introduced into the RMI compass rose.

6. At Figure 4-2 the compass rose now underreads by 10°. The ADF needle still shows 090°relative but the indicated bearing is now 140°. The compass underreads by 10°, so does the ADFbearing on the RMI.

7. The VOR needle continues to show the correct bearing of 270° to the station.

8. The following examples illustrate the relationship between the RMI and the position of theaircraft relative to the VOR/NDB stations.





EXAMPLE 4-1EXAMPLE

An aircraft is tracking towards VOR A maintaining the 140° radial with 17° of port drift. NDB Bbears 220° relative to the aircraft. VOR A and NDB B are approximately equidistant from theaircraft.

(a) Draw a diagram to show the position of the aircraft, the VOR and NDB.

(b) Draw a diagram to show the appearance of an RMI on which both VOR and ADFbearing information is shown.





SOLUTION

FIGURE 4-3





EXAMPLE 4-2EXAMPLE

An aircraft is tracking away from VOR C maintaining the 245° radial with 15° of starboard drift.NDB D lies to the west of VOR C. Heading will be altered to track inbound to NDB D when NDBD is abeam track, the drift will then be 10° port.

Draw a diagram to show the aircraft RMI on which both ADF and VOR information is presented:

(a) shortly before turning

(b) shortly after turning.

SOLUTION

See Figure 4-4 and Figure 4-5. Figure 4-4 is not required as part of the answer, but it is obviouslymuch easier to see what's what with the aid of such a diagram.





FIGURE 4-4

FIGURE 4-5





In-flight ADF and VOR Procedures using an RMI9. The following procedures, as detailed in ICAO Doc 8168, are those to be adopted when usingan RMI in conjunction with ADF/VOR equipment in the air.

Homing to an NDB/VOR Beacon10. In a pure homing situation the aircraft heading is adjusted to keep the appropriate RMIpointer (ADF or VOR) aligned with the heading index, i.e. the nose of the aircraft is always pointeddirectly towards the beacon. Unless the wind is ‘light and variable’ the aircraft heading will have tobe continually changed during the homing, with the aircraft flying a curved path to the station,ultimately arriving overhead the beacon but facing into wind.

Tracking to an NDB/VOR11. The ‘homing’ procedure, as described above, is generally inappropriate in the commercialworld as the overriding requirement is usually to maintain a given track (e.g. maintaining the centre-line of an airway). Say, for example, that an aircraft is required to track inbound to a VOR/NDB ona track (QDM) of 090°M and at the same time it is experiencing 20° starboard drift. Assuming theaircraft is currently on track, the heading should be adjusted to 070°M and the RMI will indicate abearing of 090°M to the VOR/NDB (see Figure 4-6). If the drift were 15° port the heading wouldhave been 105°M with the RMI still indicating a bearing of 090°M to the beacon. In both of theseexamples, providing the drift remains constant, all indications will remain the same and the aircraftwill maintain track until reaching the overhead.





FIGURE 4-6Tracking using an RMI

12. If during the procedure described above the RMI reading starts to increase, the aircraft will bedrifting port of track and it will be necessary to allow for less drift than was originally anticipated. Ifthe RMI reading starts to decrease, the aircraft will be drifting to starboard of track and therefore itwill be necessary to allow for more than 20° starboard drift.





Tracking from an NDB/VOR13. Imagine that in Figure 4-6 the aircraft has now flown overhead the beacon and is required totrack outbound on the same track. Assuming the aircraft is on track initially, and the drift is still20°S, then by flying the same heading of 070°M the aircraft will remain on track with the RMI nowreading 270°M. If the RMI reading changes then the heading should be adjusted in the same manneras before (i.e. if the RMI reading increases the aircraft will be drifting starboard of track; if thereadings are decreasing the aircraft will be drifting to port).

Interceptions14. If, in the above tracking procedures, the aircraft becomes a long way off-track it may benecessary to re-establish yourself by intercepting the particular track again. For example, inFigure 4-7, if the aircraft is several miles to starboard of the inbound track (360°M) to the NDB/VOR, then an intercept heading should be selected (normally about 30° to the desired track - in thiscase say, 330°M), and the aircraft flown until the RMI pointer is just approaching 360°M. Theaircraft would now be turned onto a heading of 360°M ± drift allowance.





FIGURE 4-7Interceptions using an RMI

15. In practice the turn onto the required track is made with a few degrees still ‘to go’ on theRMI. This angle of lead allows for the fact that the aircraft cannot turn on a spot and will thereforeuse up some distance in the turn. Providing the angle of lead is chosen correctly, bearing in mind therange from the ground beacon, the aircraft should roll out onto the correct track, on the correctheading.





Procedure Turns16. A procedure turn, as illustrated at Figure 4-8, is a commonly used method of reversing thetrack of an aircraft during say, a non-precision approach, whilst ensuring that the aircraft stayswithin a safe sector. The procedure turn itself is a timed procedure, usually involving a 45 sec (stillair) straight leg, along a predetermined track (150°M in Figure 4-8), however the RMI may be usedto facilitate accurate track keeping on the outbound and inbound parts of the procedure (either sideof the procedure turn).

FIGURE 4-8Procedure Turns using an RMI

17. In Figure 4-8 the RMI would be used in the manner described in paragraph 13 to follow theoutbound track of 105°M. Having almost completed the procedure turn, the RMI would then beused to intercept a track of 285°M inbound to the beacon in the manner described in paragraph 14.





Holding Patterns18. Holding patterns are race-track type patterns with all turns in the pattern being either right-hand (RH pattern) or left-hand (LH pattern). The procedure is indicated on terminal approach charts(TAPs). To enter the pattern, depending on the direction of approach to the VOR/NDB it may benecessary to carry out a pre-entry manoeuvre. 360° approach directions around the beacon aredivided into three approach sectors as shown in Figure 4-9. Each sector has its own specific arrivalprocedure.





FIGURE 4-9Entry Procedure for RH and LH Holding Patterns





19. To clarify the sector division, review the RH holding pattern in Figure 4-9. The inbound trackis 270°(M). The sector divisions based on this track are as follows :

20. If your approach track is along one of the dividing lines you may choose either sector. Thedetails of each sector join are as follows :

Sector 1

(a) On arrival overhead, fly parallel to the reciprocal of the inbound leg for theappropriate time.

(b) Turn left and home back to the NDB/VOR.

(c) On second arrival over the facility turn right and commence the pattern.

Sector 2

(a) On arrival overhead, make good a track 30° to the reciprocal of the inbound legtowards the inside of the pattern.

(b) At the appropriate time, turn right and join the pattern on the inbound leg, and hometo the facility.

Sector 3

(a) On arrival overhead, join the pattern directly.

Sector 1 - approaching the beacon between 090°(M) and 200°(M)







21. In each case the holding pattern proper, commences when overhead the beacon. Turn rightthrough 180° rate one and start timing when abeam the beacon. In zero wind conditions an RMIpointer indicating 090° relative to the nose of the aircraft, (or 180°M on the RMI) indicates theabeam position. To this figure add the amount of drift if it is starboard, subtract from it if it is port.For example in Figure 4-10 the aircraft has 20°S drift and therefore the outbound timing would startas the relative bearing indicates 110° (i.e. 180°M again on the RMI).

FIGURE 4-10Holding Pattern Procedures

22. The outbound track is parallel to the inbound track. Apply drift and adjust the leg timing forground speed. At the end of the outbound leg, turn right through 180°, rate one and fly inboundusing the RMI to track accurately to the VOR/NDB.

23. A LH race-track pattern is flown in a similar manner to that described above.





NOTE:

Now read Chapters 8 and 9







DME Frequencies

Interrogation Rates

Beacon Saturation

Maximum Range Limitations

Additional DME Functions

DME Accuracy

VOR-DME Frequency Pairing

Airborne Equipment

System Integrity

Use of a DME to Fly a DME Arc

DME Chart Symbology

Tacan





5Distance Measuring Equipment

1. Distance measuring equipment (DME) is a pulsed secondary radar system. The main purposeof the equipment is to display to the pilot the range of his aircraft from a fixed ground station.

Principle of Operation2. A series of double pulses, or pulse pairs, is transmitted in all directions by the aircraft DMEequipment. Providing that the aircraft is within line of sight range of the ground DME station (thetransponder) the pulsed energy from the aircraft will be received, amplified and re-transmitted by thetransponder, again in all directions. The pulse train which has been re-transmitted by thetransponder will arrive back at the aircraft DME receiver. The airborne equipment measures thetime interval between transmission and reception of each pulse pair and, using this time interval andthe known constant speed of propagation, determines the slant range of the aircraft from thetransponder. The pulses travelling from the aircraft to the transponder on the ground are termed theinterrogation pulses, and the identical sequence of pulses travelling from the transponder to theaircraft the response pulses.

DME Frequencies3. DME operates in the UHF band between 960 and 1213 MHz. Frequency spacing is at 1MHz intervals. A DME channel consists of two carrier wave frequencies, always 63 MHz apart.For example DME channel 1 uses a carrier wave frequency of 1025 MHz for the interrogation pulsetrain, and a carrier wave frequency of 962 MHz for the response pulse train.





4. The advantages of secondary radar over primary radar are discussed in chapter 8. One of theadvantages of secondary radar is that the pulse trains can be coded. With DME the pulse trains arenot in fact coded (but they are unique, as will be explained shortly). The other advantage ofsecondary radar, namely that equipment is smaller and lighter, most certainly does apply to DME.The technique of changing the carrier wave frequency at the transponder is made necessary withDME for the following reasons:

(a) The aircraft DME transmitter fires a continuous series of pulse pairs in all directions.Let us take just one of these pulse pairs and consider what happens. The pulse pairtravels to the DME transponder on the ground, is regenerated, and fired back to theaircraft. Unfortunately the aircraft DME receiver will already have received thereflection of the interrogation pulse pair either from the ground immediately belowthe aircraft or from nearer heavy clouds (weather clutter). In order to enable theaircraft DME receiver to distinguish between ground (or weather) reflected pulses andthe reply pulses from the transponder, the carrier wave frequency from thetransponder is always 63 MHz above or below the airborne transmitter frequency.

(b) If no change of frequency were made at the transponder, the transponder itself couldbecome self-triggering. This would occur whenever a response pulse from the groundstation bounced off a nearby surface or cloud and returned to the transponder, to allintents and purposes as a further interrogation pulse.





Interrogation Rates5. When a DME channel is initially selected, the airborne equipment searches for a range lock-on. During this search period, the airborne equipment interrogates the transponder, initially, at therate of approximately 150 pulse pairs per second. In order to avoid beacon saturation, if lock-on hasnot been achieved once the airborne equipment has transmitted 15,000 pulse pairs (approximately100 seconds), the interrogation rate is reduced to 60 pulse pairs per second.

6. Once the airborne equipment has locked-on to the reply pulses from the transponder, theinterrogation rate decreases to approximately 24 pulse pairs per second. The term lock-on simplyindicates that the airborne equipment has established the aircraft's slant range from the groundstation. In the event of the temporary loss of an incoming signal at the aircraft, for example duringan aircraft turn when a wing may well interrupt the line of sight signal path, the airborne equipmentwill be prevented from commencing another range search for a period of 10 seconds.

Beacon Saturation7. The ground transponder is capable of transmitting only 2700 pulse pairs per second.Accepting that the majority of aircraft using a given transponder will be locked-on (24 pulse pairsper second), whilst a few will be in the range search mode (150 or 60 pulse pairs per second) it isreasonable to assume that the average number of pulse pairs per aircraft per second is 27. It cantherefore be seen that the transponder can only provide range information to 100 aircraft or so. Ifmore than this number of aircraft interrogate a single transponder, it is said to be saturated. Underthese circumstances, the ground beacon will reply only to the strongest signals, which are likely tohave originated from the nearest aircraft.





Maximum Range Limitations8. DME operates in the UHF band and is therefore limited to line of sight range. Maximumtheoretical range may be determined for an aircraft at a given altitude using the by now familiarformula:

It is for this reason that DME range display indicators generally do not exceed 300 nm (the maxrange at which an aircraft at approximately 58,000 ft would receive a ground beacon whoseelevation is MSL).

Additional DME Functions9. By monitoring the rate of change of range, the airborne DME equipment is able to determinethe aircraft's groundspeed. A read-out of groundspeed derived by DME will be reasonably accurateonly if the aircraft is tracking directly towards, or directly away from, the transponder. Even withthe aircraft tracking directly over the transponder, the groundspeed read-out will not be totallyaccurate, especially at close range, since the equipment is calculating groundspeed using rate ofchange of slant range rather than horizontal range.

10. By integrating slant range and groundspeed, the equipment is also capable of calculating timeto station (TTS). Obviously this figure will only be acceptably accurate if the aircraft is trackingdirectly towards the ground station.

Maximum Range(nm) 1.25 H1 H2+( )=





DME Accuracy11. Considering solely the accuracy of the slant range read-out, DME is inherently very accurate.The slant range may be considered to be accurate to within ±¼nm plus 1.25% of range; this being anICAO requirement. There is a small time delay at the transponder between the reception and the re-transmission of each pulse. This delay (of 50 microseconds) is called the Echo Protection Circuit andcaters for any reflected or echo interrogations arriving at the ground beacon a short-time after theline-of-sight interrogation. By suppressing the ground receiver for long enough after reception of theinitial interrogation the echo will not trigger a reply. The delay described is known to the airbornereceiver and is allowed for by the receiver when converting lapsed time into slant range. Rememberhowever that it is the slant range which is indicated and it is the indication of slant rather thanhorizontal range which is considered to be the major error of the equipment. Obviously, slant rangeerror is greater at short ranges - see Figure 5-1.





FIGURE 5-1Slant Range v Plan Range

VOR-DME Frequency Pairing12. VOR gives bearing information. DME gives range information. The optimum angle of cutfor a two-position line fix is a right angle, and this is achieved by using a VOR transmitter and aDME transponder which are situated at the same point on the ground.

13. To simplify matters VOR frequencies and DME channels are paired, such that each VORfrequency has a DME channel assigned to it under an ICAO agreement. There are 126 DME ‘X’channels available (the ‘X’ notation is explained shortly) and those which remain spare when eachVOR frequency has been paired are allocated to ILS localiser frequencies (in order to give range fromthe touchdown point during an approach) and to some VHF R/T frequencies.





14. The ‘X’ notation following the appropriate DME channel number denotes that theinterrogation carrier wave frequency is 63 MHz higher than the response carrier wave frequency forchannels 1 to 63 and 63 MHZ lower than the response carrier wave frequency for channels 64 to126.

15. With the advent of additional VOR frequencies, such as Clacton which transmits on 114.55MHz, there are now a greater number of VOR frequencies than DME ‘X’ channels, and to overcomethis shortfall a further 126 DME ‘Y’ channels have been introduced, simply by reversing therelationship between DME interrogation and response carrier wave frequencies. You won't find achannel X/Y selector on the DME control panel, the equipment automatically takes care of thisselection. This is achieved by the spacing of the transmitted pulse pairs. In ‘X’ channel transmissionsthe leading edges of the two pulses in each pair are 12 microseconds apart; in ‘Y’ transmissions theyare 36 microseconds apart. The airborne transmitting equipment is programmed to select theappropriate pulse spacing.

16. As an actual example of the use of X and Y channels in the frequency pairing of VORs andDMEs, consider the following :

(a) A VOR on a frequency of 112.30 MHz is always paired with a DME on channel 70X(interrogation frequency 1094 MHz ; response frequency 1157 MHz).

(b) A VOR on a frequency of 112.35 MHz is always paired with a DME on channel 70Y(interrogation frequency 1094 MHz ; response frequency 1031 MHz).

17. In order to establish clearly the relationship between a VOR station and any DME stationoperating on the paired channel, the Morse identifiers of both should be carefully checked. The tableat Figure 5-2 outlines the information which can be gleaned from a sensible comparison of the twoidentifiers. First note the following points:





The term ‘associated’ which is used at Figure 5-2 denotes that the VOR transmitter and the DMEtransponder are either:

(a) co-located, or

(b) situated so close to each other as to make no practical difference. For en routefacilities this means that they are located within 2000 feet of each other, or forterminal approach aids within 100 feet of each other.

The DME transmits its identifier once every 30 seconds and the VOR once every 10 seconds. Theterm ‘synchronised’ used in the right hand column of the table at Figure 5-2 indicates that each thirdVOR identifier is suppressed, and is replaced by the DME identifier. In this event the two identifiersare heard at different tones, to enable the pilot to distinguish one from the other.





FIGURE 5-2VOR / DME Frequency Pairing

Figure 5-3 shows a typical co-located conventional VOR/DME installation.

Relationship of VOR and DME Frequency Relationship

Identifiers

VOR and DME associated Paired Identical three-letter Morse groups, synchronised

VOR and DME not associated but using the same location and may be used in conjunction with each other for normal en-route (airways) navigation. The ground stations would not normally be more than 6 nm apart in this case

Paired The first two letters of the Morse identifiers are the same but the last letter of the DME identifier will be changed to a Z. Again the identifiers are synchronised

VOR and DME at entirely different locations

Unintentional frequency pairing may exist

Entirely different and unsynchronised Morse identifiers are used





FIGURE 5-3Co-located VOR/DME Installation

Airborne Equipment18. A block schematic diagram of a Boeing 737 DME system is shown at Figure 5-4.





FIGURE 5-4Boeing 737 DME Installation

19. An alternative type of airborne DME control panel is illustrated at Figure 5-5. There are twooptions for selecting DME station, one is by channel number, the other is by paired VOR frequency.With the equipment shown at Figure 5-5 the latter option is used. With the equipment shown thefunction switch gives the operator the choice of displaying either slant range (NM), time to station(MIN) or groundspeed (KT).





FIGURE 5-5Typical DME Control Panel

20. Many airborne DME sets are designed such that, by selecting the required frequency on aVOR receiver, the DME set is automatically tuned to the two carrier wave frequencies constitutingthe channel number which is paired with the VOR frequency selected. Figure 5-6 shows the controlpanel of such a DME set. The NAV 1/NAV 2 positions on the station selector allow the pilot tochoose which of the two VOR receivers normally fitted is to be coupled with the DME set. TheHOLD function enables the pilot to leave the DME set tuned to the channel associated with the VORfrequency selected on either NAV 1 or NAV 2 despite the fact that the appropriate VOR receiver issubsequently retuned to a different VOR station.





FIGURE 5-6Alternative DME Control Panel

System Integrity21. The basic principle of operation of DME has already been discussed, however furtherconsideration of the way in which the airborne equipment positively determines range is nownecessary.

22. Remember that the aircraft DME equipment fires pulse pairs which travel to the transponderand back to the aircraft at a constant speed. It is the time interval between the transmission and thereception of each pulse pair which the airborne equipment measures in order to determine range.The problem lies in the fact that the transponder may be transmitting 2700 pulse pairs every second.The airborne DME equipment will receive each and every one of these pulse pairs, despite the factthat very few of them will be responses to its own interrogation pulse train. The airborne equipmentmust therefore be able to distinguish between reply pulses to its own interrogation pulse train andunwanted reply pulses from the transponder.





23. In order to achieve this vital recognition, the airborne unit transmits a unique pulse train.That is to say, that the time between successive pulse pairs does not remain constant, but is jittered.When this unique pulse train arrives at the transponder, it is exactly reproduced and re-transmitted,albeit on a different carrier wave frequency. The airborne receiver recognises this unique pulse trainby virtue of the fact that these pulse pairs, and only these pulse pairs, will arrive back at the aircraftconsistently at a constant time interval after the interrogation pulses. To achieve this recognition, thereceiver employs strobing or time-gate circuits.

Use of a DME to Fly a DME Arc24. Certain Instrument Approach procedures require that a DME Arc be flown prior tointercepting the final approach track (i.e. the aircraft is flown such that the DME reading remainsconstant). Usually the DME is co-located with another aid, either a VOR or NDB, such that theadditional aid can be used as part of the manoeuvre.

25. Figure 5-7 shows an aircraft flying a 20 nm arc against a co-located VOR/DME. Shortlybefore 20 nm is reached, the aircraft is turned to keep the VOR reading 090° relative to the nose ofthe aircraft. By holding this particular relative bearing the DME reading should remainapproximately constant, however slight alterations of heading may have to be made to ‘fine tune’ therange indication.





FIGURE 5-7Procedure for Flying a DME Arc

DME Chart Symbology26. The following chart symbols are used in connection with a DME





FIGURE 5-8DME Chart Symbols

Tacan27. Tacan is a military version of DME, giving both range and bearing information from thetransponder. An aircraft equipped with DME may interrogate a Tacan transponder. Range,groundspeed and time to station information will be displayed in the normal manner once a lock-onis achieved, however the airborne DME equipment is not able to utilise the bearing element of theTacan system.

28. For a civilian aircraft, Tacan is tuned by selecting the VOR frequency which is paired with it(as is the case with DME). A Tacan and a VOR may serve the same location in much the same wayas a DME and a VOR. When the Tacan and the VOR ground stations are within 600 metres of eachother the same three letter identifier will be used for each. When the distance between the twostations exceeds 600 metres the final letter of the Tacan identifier is changed to a Z.






QUESTIONS:QUESTION 1.

What frequency range does DME operate within?

QUESTION 2.

There are two reasons why the interrogation and response frequencies for a particular DME channeldiffer by 63 Mhz, what are they?

QUESTION 3.

After "lock-on" an airborne DME will be interrogating a ground based transponder at a rate of:

QUESTION 4.

A DME, that is at 1600 ft AMSL, is being received by an aircraft at 36000 ft. Given that there are noother limitations, at what range would the aircraft theoretically lose the signal?

QUESTION 5.

What is the ICAO specified accuracy for a DME?

QUESTION 6.

How does a DME discriminate between X or Y channel transmissions?





QUESTION 7.

How often does a DME transmit its identification code?

QUESTION 8.

What does the term "associated" mean when used in connection with the frequency pairing of en-route VOR and DME facilities

QUESTION 9.

When using a DME with a groundspeed readout,the groundspeed calculated will be at its mostaccurate when the aircraft is:

QUESTION 10.

If a DME signal is lost, for example due to a change in aircraft attitude, how long will the equipmentrange display remain "frozen" before the equipment re-enters "search" mode?

ANSWERS:ANSWER 1.

960 to 1213 Mhz

ANSWER 2.

1. It prevents the equipment from locking-on to its own reflections

2. It prevents the ground transponder from self-triggering





ANSWER 3.

24 pulse pairs per second

ANSWER 4.

Range = 1.25 = 287nm

ANSWER 5.

± ¼nm plus 1.25% of range

ANSWER 6.

X channels have a pulse pair spacing of 12µ secY channels have a pulse pair spacing of 36 µ sec

ANSWER 7.

A DME identifies itself every 30 seconds

ANSWER 8.

The term associated means that the VOR and DME are either co-located, or are within 2000 ft (600metres) of each other

ANSWER 9.

Flying directly towards or away from the facility at long ranges

1600 36 000,+( )





ANSWER 10.

The DME memory is 10 seconds






ILS Ground Equipment

Localiser Transmitter

Localiser Radiation Pattern

Glidepath Transmitter

Glidepath Radiation Pattern

ILS Glideslope vs Visual Glideslope

Glidepath Calculations

Height and Range Calculations

Rate of Descent Calculations

ILS Frequencies

Frequency Pairing

Station Identification

ILS Calibration





The ILS Meter

Alarm Flags

Localiser Needle Sensitivity

Glidepath Needle Sensitivity

Marker Beacons

Airways Fan Markers or Z Markers

The Airborne Marker Beacon Receiver

ILS Facility Performance Categories

ILS Monitoring Stations


The ILS Approach Plate





6The Instrument Landing System

1. The Instrument Landing System (ILS) is a runway approach aid which provides the pilot withaccurate guidance both in azimuth and elevation during an approach in bad weather.

ILS Ground Equipment2. The ground installation consists of:

(a) A localiser transmitter which defines the extended centreline of the instrumentrunway, and indicates any deviation from this centreline.

(b) A glidepath transmitter which defines a safe descent slope (normally three degrees),and again indicates any deviation from this safe approach.

(c) Normally two (occasionally three) marker beacon transmitters for a typicalinstallation. That is to say that with many installations the inner marker is omitted,leaving only the middle and outer markers.

3. The primary purpose of the markers is to define specified ranges from the ILS touchdownpoint. Many modern ILS installations employ DME transponders rather than/in addition to markerbeacons to provide range information. In such an installation the DME channel is paired with theILS localiser frequency, so that in many aircraft the DME channel is automatically tuned when theILS is selected. With a paired ILS/DME installation the DME range information is zero referenced tothe ILS touchdown point. It is important to appreciate that the DME range information provided bythis system is considered to be precise only when the aircraft is in line with the runway on theapproach side. In other areas the range must be considered to be approximate.





Localiser Transmitter4. A localiser antenna array is approximately 25 metres wide and four metres high, and isnormally situated some 300 metres beyond the upwind end of the instrument runway, see Figure 6-1.

FIGURE 6-1Localiser Aerial Location

5. Should it not be possible to locate the localiser aerial on the extended centreline, it may belocated to one side of the runway, giving what is known as an offset ILS. In this case, the QDM ofthe localiser centreline will differ from the runway centreline QDM by a few degrees.

Localiser Radiation Pattern6. The localiser transmits two overlapping lobes of electro-magnetic energy (designated A8W)on the same VHF carrier wave frequency. The centre of the overlap area, the equisignal, defines theILS QDM, see Figure 6-2.





FIGURE 6-2Localiser Radiation Pattern

7. The lobe on the pilot's left during the approach is amplitude modulated at 90 Hz, whilst theright lobe is amplitude modulated at 150 Hz. The depth of modulation of both the lobes is made tovary, being greatest at the centre and least at the sides of the lobes. The airborne localiser receivercompares the depth of modulation of the 150 Hz and 90 Hz waves. When they are of equal depththe localiser needle will be centralised. When the depth of modulation is uneven the localiser needleis deflected in the appropriate direction. Note that the Difference in Depth of Modulation (DDM)increases with displacement from the centreline; hence, the greater the difference, the greater thedisplacement of the localiser needle from the centre of the instrument.





NOTE:

DDM is calculated by subtracting the percentage modulation depth of thesmaller signal from the percentage modulation depth of the larger signal, andthen dividing by 100.

8. It should be noted that on the approach side of an ILS localiser aerial there is a Course Sector,positioned equally astride the centreline of the runway, outside of which the localiser indicator willshow full scale deflection in the appropriate sense. This course sector is a maximum of 6° wide (i.e. ±3° either side of the centreline) and, within the sector, the difference in modulation depths increaseslinearly with displacement from the centreline.

9. In the United Kingdom ILS localisers which are associated with normal glidepath transmittersprovide coverage from the centre of the localiser antenna to distances of:

(a) 25 nm within plus or minus 10° of the equisignal (centre) line

(b) 17 nm between 10°and 35° from the equisignal (centre) line as illustrated atFigure 6-3.





FIGURE 6-3Localiser Coverage

10. In the United Kingdom ILS localisers which are associated with steep angle glidepathtransmitters provide coverage from the centre of the localiser antenna to distances of:

(a) 18 nm within plus or minus 10° of the equisignal (centre) line





(b) 10 nm between 10° and 35° from the equisignal (centre) line

11. As far as the above coverage areas are concerned, a normal glidepath transmitter should beconsidered to be one which produces a glidepath angle of approximately three degrees above thehorizontal, and a steep glidepath should be considered to be one which defines an angle from thehorizontal of 4° or more.

12. Pilots are warned that use of the localiser outside these areas, even on the approach side, canlead to False Course and Reverse Sense indications being received. Such use should not beattempted. In particular it must be noted that there is no provision for localiser Back Beams to beused in the United Kingdom, and any indications from them must be ignored. Tests have also shownthat FM interference, from broadcast stations transmitting on frequencies just below 108MHz, mayeffect both the localiser course guidance and alarm flag signals (see later) of the airborne installation.The effects of such interference vary depending upon the difference in depth of modulation of thelocaliser signals being processed. All modern ILS localiser receivers have an FM immune filter fittedto prevent this specific type of interference.

13. In some of the earlier ILS installations it is possible to receive false localiser signals which maybe at a considerable angle to the runway QDM. The use of low powered locator NDBs to guide theaircraft onto the correct localiser beam (see later) has helped overcome this problem.

14. In circumstances where the use of the back beam (back course) is authorised, say for example,to maintain the runway centreline outbound or following a missed approach (see paragraph 45), the‘fly left’ and ‘fly right’ demands must be reversed by the pilot in order to maintain the desired track.





Glidepath Transmitter15. There are two glidepath aerials which are both mounted on a mast approximately ten metrestall, which is displaced some 150 metres from the runway centreline and 300 metres upwind of thethreshold markings.

Glidepath Radiation Pattern16. As with the localiser, the glidepath transmitter emits two overlapping lobes of electro-magnetic energy (designated A8W) on the same carrier wave frequency. The frequency range usedfor glidepath transmissions lies in the UHF band, and in this case the lobes overlap in the verticalplane. Again the lobes are continuously amplitude modulated at 90 Hz and 150 Hz. Figure 6-4shows the idealised radiation pattern with the equisignal defining the glidepath at a typical value of3° above the horizontal plane passing through the touchdown zone.





FIGURE 6-4Glidepath Radiation Pattern





FIGURE 6-5False Glidepaths





17. Since the lower (150 Hz) lobe lies adjacent to the surface ground-reflected waves result,giving side lobes (Figure 6-5). These side lobes may produce additional equisignals and consequentlyfalse glidepaths. Fortunately, these false glidepaths will be situated above the main glidepath andcannot therefore result in an aircraft flying dangerously low during the approach should the falseglidepath be inadvertently followed. Indications that the aircraft is flying a false glidepath are listedbelow:

(a) During a normal ILS procedure, the aircraft captures the glidepath from below. Thisbeing the case, the true glidepath (being the lowest) will be the first one to beintercepted. The Civil Aviation Authority has issued a warning to pilots emphasisingthat special care must be taken at certain airfields around the world where proceduresare published involving capture of the glidepath from above.

(b) The first (lowest) false glidepath will give a descent slope which is inclined at least 6°to the horizontal for a normal 3° glidepath, or 5° to the horizontal for a 2.5°glidepath. This will result in a rate of descent of at least twice the expected value.

(c) The approach plate used by the pilot (see Figure 6-17) during an ILS approach showscheck heights and altitudes at the marker beacons, and locator beacons if appropriate.If a false glidepath has been captured, a check of the altimeter will verify this. Atypical check height over the outer marker would be 1500 feet (QFE), whereas on thefirst false glidepath the altimeter would read 3000 feet (QFE) or above.

18. Glidepath coverage in azimuth (for United Kingdom installations) is provided through an arcof 8° on either side of the localiser centreline out to a range of 10 nm from the threshold, asillustrated at Figure 6-6.





FIGURE 6-6Glidepath Coverage in Azimuth

19. For glidepath transmitters which produce a steep glidepath, the coverage is reduced to a rangeof 8 nm from the threshold, again through an arc of 8° on either side of the localiser centreline.

20. Glidepath coverage in elevation is provided through an arc of 1.35° above the horizontal to5.25° above the horizontal. These figures apply to a standard 3° glidepath installation, and are basedon the formulae which state that glidepath coverage (in elevation) is provided through an arc(measured from the horizontal) of between glidepath angle x 0.45 and glidepath angle x 1.75, asillustrated at Figure 6-7.





FIGURE 6-7Glidepath Coverage in Elevation





21. Pilots are warned that use of the glidepath outside these limits can lead to intermittent andincorrect indications being received. In particular, use of the glidepath at very shallow approachangles, (that is below 1500 ft aal at 10 nm range), should only be attempted when the promulgatedglidepath intercept procedure requires such use.

22. The glidepath indication must be ignored if the approach angle is so shallow as to put theaircraft at a height of 1000 ft or below at a range from touchdown of 10 nm or more.

23. Certain glidepaths in the United Kingdom do not exhibit correct deflection sensitivity to oneside of the localiser course line. This effect is caused by terrain or other problems and can lead toinadequate fly up indications being received. When this situation exists a warning will bepromulgated by NOTAM and subsequently appear in the appropriate columns of the COM 2 sectionof the UK AIP.

ILS Glideslope vs Visual Glideslope24. Where no obstacle clearance problem exists at an airfield both the ILS glideslope and visualglideslope (either Precision Approach Path Indicator [PAPI] or Visual Approach Slope Indicator[VASI]) will normally be set at around 3° (check the relevant documentation) and therefore thereshould be direct correlation between the information provided by both systems. Unfortunately on along bodied aeroplane, such as a Boeing 747 or A300 Airbus, the wheels of the aircraft will be muchfurther below the pilots eyes and it is important that his eyes follow a parallel, but higher slope, toensure adequate wheel clearance at the runway threshold. To facilitate this, a 3 bar VASI has beendeveloped such that pilots of Long Bodied aircraft use only the second and third wing bars of theVASI and ignore the first bar (the lower one). When on the correct visual glideslope the top bar willappear red and the middle bar white.





Glidepath Calculations25. Problems involving calculation of aircraft height at a given range from the threshold, or ofexpected rate of descent, are frequently set in the examination.

26. There are essentially three methods for solving these types of problems. Firstly the rule ofthumb method based on the 1 in 60 rule but assuming a 6000 ft nautical mile. This is ideal for grosserror checks when flying, since you can do it in your head (the height reduction per nm on a 3°glidepath is 300 ft, on a 3.5° glidepath is 350 ft, and so on). The second option is the 1 in 60 rulemethod using a 6080 ft nm, the third option involves basic trigonometry, and both of these methodsare considered in the following examples.





Height and Range CalculationsEXAMPLE 6-1

EXAMPLE

Determine the height above touchdown of an aircraft which is on a 3° glidepath at a range of 3 nmfrom the threshold.

SOLUTION

FIGURE 6-8

Figure 6-8 illustrates the solution, which may be formulated in either of two ways:

(a) Using the 1 in 60 rule

[range from threshold (in ft) + 1000 ft] glidepath angle160------×× height (ft)=





EXAMPLE 6-2

This will give a height above touchdown of:

or alternatively, using trigonometry:

This will give a height above touchdown of:

3 6080×( ) 1000+[ ] 3×60

---------------------------------------------------------- 962 ft=

Tan (glidepath angle) range from threshold in ft 1000+( )× height (ft)=

Tan 3° 3 6080×( ) 1000+[ ]× 1000ft=

EXAMPLE

Determine the height above touchdown of an aircraft which is on a 3.25° glidepath at a range of3.75 nm from the runway threshold.

SOLUTION

Using the 1 in 60 rule:

Using trigonometry:

The tangent of 3.25° is 0.057

[(3.75 6080 ) 1000 ] 3.25×+×60

--------------------------------------------------------------------------- 1289 ft=

Tan 3.25° 3.75 6080×( ) 1000+[ ]× 1357 ft=





27. It is important to appreciate that in both the previous examples the range has been given fromthe runway threshold, and accordingly the 1000 ft distance to the ILS touchdown has been includedin the range calculation. If, in the question, the range is given from the ILS touchdown point, the1000 ft must be omitted from the calculation.

Rate of Descent Calculations28. A similar procedure is necessary in the examination when calculating the aircraft's rate ofdescent for a given glideslope and groundspeed.





EXAMPLE 6-3EXAMPLE

Determine the rate of descent required for an aircraft on a 3° glideslope at a groundspeed of 175 kt.

SOLUTION

Using the 1 in 60 rule:

Using trigonometry:

=

=

=

Rate of descent (ft/m in) Groundspeed (in ft/min) glideslope

160------××

(175 6080 ) 3××60

-----------------------------------------160------×

887 ft/min

=

=

=

Rate of descent (ft/m in) Groundspeed in ft/min Tan (glideslope)×

175 6080×60

--------------------------- Tan 3°×

922 ft/min





ILS Frequencies29. Two bands of frequencies are allocated solely for use with ILS installations.

Localiser30. ILS localisers transmit on one of forty allocated frequencies in the VHF band. The frequencyrange is from 108.1 MHz to 111.95 MHz at 50 KHz (0.05 MHz) channel spacing, (e.g. 108.1,108.15, 108.3, 108.35 to 111.9 and 111.95 MHz but not 108.2, 108.25 and other frequencies withan even number following the decimal point, since these frequencies are used by VOR).

Glidepath31. ILS glidepath transmissions are in the UHF band, again with 40 frequency channels rangingfrom 329.15 MHz to 335 MHz, with 150 KHz (0.15 MHz) channel spacing.

Frequency Pairing32. Localiser and glidepath transmissions are always frequency-paired. Consequently eachlocaliser frequency has a glidepath frequency associated with it. When the pilot selects a localiserfrequency on the VOR/ILS control unit, the glidepath receiver is automatically tuned to the correctglidepath frequency.

33. The advantages of frequency pairing are:

(a) Flight deck workload is reduced.





(b) Safety. Excluding equipment malfunction, it is impossible to select the wrongglidepath frequency for a given localiser frequency.

Station Identification34. Station indentification is achieved by modulating onto the localiser carrier wave ahorizontally-polarised, 1020Hz amplitude modulated tone to give the ident. The ident may be two ormore letters in morse transmitted at a rate of six or more words per minute. If it is necessary todistinguish an ILS quickly from other navigation aids, the ident may be preceded by the letter ‘I’. Insome category 1 or 2 ILS systems, ground-to-air communications (i.e. voice) may be superimposedonto the localiser carrier providing that it doesn’t interfere with the ident or normal localiseroperation.

35. When an ILS is undergoing maintenance, or is radiating for test purposes only, the identifierwill either be removed completely or replaced by a continuous tone. Under these conditions noattempt should be made to use the ILS as completely erroneous indications may be received.

36. Additionally, in some instances, because of an unserviceable glidepath, the ILS may beradiating for localiser approaches only. In this case the identification coding will continue to beradiated, and a warning to the effect that the ILS is radiating for localiser only approaches will begiven by the ATC. In this situation (localiser only approach) the glidepath may be radiating forsetting up purposes or for flight inspection. UNDER NO CIRCUMSTANCES should the glidepathbe used at this time, the glidepath radiation pattern may be subject to interruptions and alterationswithout warning, and may at all times be given erroneous indications.





ILS Calibration37. All ILS systems in the UK are regularly checked for accuracy (calibrated) by aircraft equippedwith a sophisticated airborne monitoring system. The localiser is checked for accuracy out to a rangeof 10 nm, and is further checked to ensure that it is free from interference out to a range of 25 nmand to a height of 6250 feet.

38. The glidepath is checked for accuracy out to a range of 10 nm through a horizontal arcextending through eight degrees on either side of the localiser centreline.

The ILS Meter39. Many modern aircraft are fitted with horizontal situation indicators (HSI's) whichincorporate a localiser beam bar and glideslope pointer. However, the following paragraphs dealwith the basic ILS meter, which, depending on the frequency selected, serves a dual role as eitherVOR omni-bearing indicator (the horizontal needle being inoperative), or ILS meter.

40. Figure 6-9 shows a typical ILS/VOR meter. The instrument illustrated is known as a five-dotdisplay, other types use a different number of dots (often four); the ring at the centre of the displayalways constitutes the first dot.





FIGURE 6-9Typical ILS / VOR Meter

41. The OBS window has been omitted from Figure 6-9 so as to emphasise that the omni-bearingselector serves no control function when the instrument is used in the ILS mode. Many pilots elect todial the ILS QDM into the window, this is done purely as a reference and will in no way affect thelocaliser needle indications.

Alarm Flags42. The ILS system is designed to enable the aircraft to fly in close proximity to the ground in badweather. It is therefore essential that prominent alarm flags be fitted to the ILS meter to giveimmediate indication of equipment unreliability, see Figure 6-9.

43. Either or both alarm flags will become visible under any the following circumstances:





(a) Following any significant distortion of the ground transmitter radiation patterns.

(b) Should the aircraft fly outside the ILS service area - that is, outside the radiationpattern of either the glidepath or localiser transmitter. Note that the service area isNOT confined to the calibrated coverage.

(c) Following failure of either ground or airborne equipment.

(d) Following intentional or inadvertent switching off of either ground or airborneequipment or following power failure.

Localiser Needle Sensitivity44. When used in the VOR mode, the vertical needle has a sensitivity of two degrees per dot ofdeviation. When used in the ILS mode, the needle is far more responsive, and has a sensitivity of onedot for each half-degree of deviation from the localiser. As with the VOR, remember that the needlerepresents the required track, and the centre of the instrument the aircraft. At Figure 6-9 the needleis showing a fly left indication, the aircraft being 1.0 degree to the right of the centreline.

45. When executing a procedural ILS pattern it may be necessary to fly outbound along thelocaliser, see Figure 6-17. You must appreciate that the basic ILS meter always assumes that theaircraft is inbound and the sense indications (fly right/fly left) are presented accordingly. Thus, foran aircraft outbound, keeping the localiser on its right, the localiser needle will show fly left, seeFigure 6-10.





FIGURE 6-10Localiser Indications

Glidepath Needle Sensitivity46. With modern ILS installations it is safe to assume a reasonably linear rate of needle deviationwith vertical angular displacement from the glidepath. Full-scale deflection of the needle will occurwhen the aircraft is displaced by approximately 0.75 degrees above or below the glidepath. Taking astandard five-dot display, one dot displacement of the needle represents approximately 0.15 degreesof deviation above or below the glidepath.





47. As with the localiser, remember that the needle represents the glidepath and the centre of theinstrument, the aircraft. Thus at Figure 6-9, the needle is showing a fly up indication, the aircraftbeing 0.375° below the glidepath.

48. For safety reasons, half full-scale fly up indication is considered to represent the maximumsafe deviation below the glidepath. Half full-scale deflection will of course be 2.5 dots on a five-dotdisplay and this maximum safe fly up indication is shown at Figure 6-9.

Marker Beacons49. Marker beacons radiate fan-shaped patterns of energy vertically upwards. Figure 6-11 showsan installation using three marker beacons, although the inner marker is not often used these days.All marker beacons transmit on a set frequency of 75 MHz. Notice from Figure 6-1 and Figure 6-11that there is no interference between adjacent beacons because of the narrow extent of the radiationpatterns along the glidepath.





FIGURE 6-11Marker Beacon Installation





50. The marker beacon transmissions are amplitude modulated with dots and/or dashes at giventones. As the aircraft flies through the radiation pattern associated with a given marker beacon, thepilot will receive both aural and visual indications as described at Figure 6-12.

51. One or two locator beacons (low powered NDBs) are often positioned at the same sites at theOuter Marker and Middle Marker (if only one locator is used it is usually co-located with the OuterMarker). The purpose of the locators is to assist the pilot when tracking to the station or whenjoining the ILS procedure, to provide a holding facility, and to provide a cross check when passingover the markers. In addition, the benefit of having a fan marker at the locator position is that it canbe used to determine a relatively accurate overhead whilst the aircraft is flying through the locatorscone of silence.





FIGURE 6-12Marker Beacon. Aural and Visual Indications.

Airways Fan Markers or Z Markers52. Marker beacons are still sometimes found straddling airway centrelines to denote reportingpoints. As with the ILS marker beacons, the airways fan markers radiate a fan-shaped pattern on afixed carrier wave frequency of 75 MHz, however the power transmitted by an airways marker isconsiderably greater to facilitate high altitude reception. The aural identifier is a single Morse letterof high pitch tone (3000 Hz) which activates the white (inner marker) light on the aircraft markerbeacon panel.

53. In addition to the fan markers previously described there are also Z Markers which radiateenergy in a vertical cone-shaped pattern. Since all marker beacons radiate energy predominantlyupwards, it is impossible to home towards them. Markers therefore serve the sole function ofproviding a range or position check.

Outer Marker

Aural:

Visual:

Low pitch (400 Hz) dashes

A blue light flashing in synchronisation with the audible dashes at the rate of two per second

Middle Marker

Aural:

Visual:

Medium pitch (1300 Hz) alternate dots and dashes

An amber light flashing in synchronisation with the audible dots and dashes at the rate of three characters per second

Inner Marker

Aural:

Visual:

High pitch (3000 Hz) dots

A white light flashing in synchronisation with the audible dots at the rate of six per second





The Airborne Marker Beacon Receiver54. The aircraft marker beacon receiver is normally automatically switched on when the VOR/ILS is in use. A typical marker beacon panel is illustrated in Figure 6-13. The high/low sensitivityswitch positions govern the receiver gain and also the brightness of the flashing lights. The lightsshould always be checked by depressing the switch to the test position (which illuminates all threelights simultaneously) before commencing an ILS approach.

FIGURE 6-13Marker Beacon Display Panel

55. The current trend seems to be to replace ILS marker beacons with a single DME transponder,thus providing the pilot with a continuous range from the ILS touchdown point throughout theapproach.





ILS Facility Performance Categories56. ILS is used to provide guidance down to the pilot's decision height. If, at the decision height,the pilot does not have the specified visual references required to continue the approach and landingvisually, missed approach action must be initiated.

57. The decision height depends in part on the ICAO performance category of the groundinstallation. There are three categories which are defined at Figure 6-14. It will be appreciated bylooking at Figure 6-14 that, as the category of ILS increases, the lower the permissable decisionheight; hence the ILS equipment accuracy requirements will need to be much greater.

FIGURE 6-14ILS Categories

ILS

CAT

Accurate Guidance Provided Down To Decision Height RVR

1 A height of 200 ft above the horizontal plane containing the runway threshold

Not lower than 200 ft Not less than 550 m

2 A height of 50 ft above the horizontal plane containing the runway threshold

Lower than 200 ft

but

Not lower than 100 ft

Not less than 300 m

3A and along the runway (if any)

Lower than 100 ft

Not less than 200 m





58. Please appreciate that category 2 and 3 approaches may only be flown when all of theappropriate requirements are met. These include aircraft equipment, ground equipment, aerodromeprocedures and flight crew qualification and training.

ILS Monitoring Stations59. Both localiser and glidepath transmitters are automatically monitored by monitoringequipment located in an area of guaranteed reception within the normal service sector. The groundmonitor station will check for the following:

(a) a localiser shift of more than 35 ft from the centreline.

(b) a glideslope angle change of more than 0.075 x basic glidepath angle.

(c) a reduction in power output of 50% or more of any of the transmitters.

3B and along the runway (if any)

Lower than 50 ft

Less than 200 m but not less than 75m

3C along the runway and to the parking bay

no decision height or RVR limitations

Note. It must be appreciated that the above table quotes the minimum decision heights for each category of ILS. Actual decision heights at specific airfields may be higher, because of factors such as surrounding topography (which will affect the terrain clearance during a missed approach procedure), aircraft equipment, pilot currency on type and so on.





60. In any of the above circumstances, the monitoring unit will provide warning to a designatedcontrol point and cause any of the following to occur before a standby transmitter is brought intouse:

(a) the cessation of all radiations

(b) the removal of the ident signal and/or the navigational information (i.e. localiser andglide path)

(c) if the ILS is category 2 or 3, the monitor may permit operation to a lower category, i.e.1 or 2.

Factors Affecting Range and Accuracy61. The accuracy of the guidance information provided by an ILS equipment is dependent onseveral factors, some of the most important of which are discussed in the following paragraphs.

Beam Bends62. Local terrain can have the effect of bending localiser beams at some airfields and pilots will berequired to make a small, but nevertheless noticable, heading change to maintain the centreline. Thelocaliser indications will, of course, become more accurate as the aircraft approaches the runwaythreshold.

Scalloping63. The problem of propogation at ‘long’ ranges over an even ground, and when the aircraft is atlow altitude, were discussed fully in paragraph 21 and paragraph 22.





Beam Noise64. Under certain conditions the localiser and glideslope indications will fluctuate over a shortperiod of time; the name given to these fluctuations is beam noise. When beam noise is experiencedduring an ILS approach the situation must be monitored carefully, if necessary by using informationfrom alternative sources.

Sensitive and Critical Areas65. Interference to ILS signals is dependant on the total environment around the ILS antennas,and the antenna characteristics. Any large reflecting objects, including vehicles or fixed objects suchas structures within the radiated signal coverage, will potentially cause multipath interference withthe ILS localiser and glidepath signals. The location and size of the reflecting fixed objects andstructures in conjunction with the directional qualities of the antennas will determine the course orglideslope quality whether Category I, II or III. Moveable objects can degrade this structure to theextent that it becomes unacceptable. The areas within which this degradable interference is possible,need to be defined and recognised. For the purposes of developing protective zones, these areas aredivided into two types, i.e. critical areas and sensitive areas:

(a) The ILS critical area is an area of defined dimensions about the localiser and glidepathantennas where vehicles, including aircraft, are excluded during all ILS operations.The critical area is protected because the presence of vehicles and/or aircraft inside itsboundaries will cause unacceptable disturbance to the ILS signal-in-space.





(b) The ILS sensitive area is an area extending beyond the critical area where the parkingand/or movement of vehicles, including aircraft, is controlled to prevent the possibilityof unacceptable interference to the ILS signal during ILS operations. The sensitive areais protected against interference caused by large moving objects outside the criticalarea but still normally within the airfield boundary.

66. Typical examples of critical and sensitive areas that need to be protected are shown inFigure 6-15 and Figure 6-16. To protect the critical area, it is necessary to normally prohibit all entryof vehicles and the taxiing or parking of aircraft within this area during all ILS operations. Thecritical area determined for each localiser and glidepath should be clearly designated. Suitable signaldevices may need to be provided at taxiways and roadways which penetrate the critical area torestrict the entry of vehicles and aircraft. With respect to sensitive areas, it may be necessary toexclude some or all moving traffic depending on interference potential and category of operation. Itwould be advisable to have the aerodrome boundaries include all the sensitive areas so that adequatecontrol can be exercised over all moving traffic to prevent unacceptable interference to the ILSsignals. If these areas fall outside the aerodrome boundaries, it is essential that the co-operation ofthe appropriate authorities be obtained to ensure adequate control. Operational procedures need tobe developed for the protection of sensitive areas.





FIGURE 6-15Typical Localiser Critical and Sensitive Area Dimension Variations for a 3,000 m (10,000 ft) Runway





FIGURE 6-16Typical Glidepath Critical and Sensitive Area Dimension Variations





67. The size of the sensitive area depends on a number of factors including the type if ILSantenna, the topography, and the size and orientation of man-made objects, including large aircraftand vehicles. Modern designs of localiser and glide path antennas can be very effective in reducingthe disturbance possibilities and hence the extent of the sensitive areas. Because of the greaterpotential of the larger types of aircraft for disturbing ILS signals, the sensitive areas for these aircraftextend a considerable distance beyond the critical areas. The problem is aggravated by increasedtraffic density on the ground.

68. In the case of the localiser, any large objects illuminated by the main directional radiation ofthe antenna must be considered as possible sources of unacceptable signal interference. This willinclude aircraft on the runway and on some taxiways. The dimensions of the sensitive areas requiredto protect Category I, II and III operations will vary, the largest being required for Category III. Onlythe least disturbance can be tolerated for Category III, but an out-of-tolerance course along therunway surface would have no effect on Category I or II operations.

69. In the case of the glidepath, experience has shown that any object penetrating a surface abovethe reflection plane of the glidepath antenna and within azimuth coverage of the antenna must beconsidered as a source of signal interference. The angle of the surface above the horizontal plane ofthe antenna is dependant on the type of glidepath antenna array in use at the time. Very largeaircraft, when parked or taxiing within several thousand feet of the glidepath antenna and directlybetween it and the approach path, will usually cause serious disturbance to the glidepath signal. Onthe other hand, the effect of small aircraft beyond a few hundred feet of the glidepath antenna hasbeen shown to be negligible.





70. Experience has shown that the major features affecting the reflection and diffraction of theILS signal to produce multipath interference are the height and orientation of the vertical surfaces ofaircraft and vehicles. The maximum height of vertical surface likely to be encountered must beestablished, together with the ‘worst case’ orientation. This is because certain orientations can causeout-of-tolerance localiser or glidepath deviations at greater distances than parallel or perpendicularorientations.

NOTE:

The following are factors which affect the size and shape of the critical andsensitive areas : - aircraft types likely to cause interference- antenna aperture- antenna type (log periodic dipole/dipole, etc)- type of clearance (single/dual frequency)- category of operations proposed- runway length- static bends

71. When protection to the degree required for category two or three operation are in force,pilots of arriving and departing aircraft will be informed that low visibility operations (LVOs) are inforce.





72. Bearing in mind the above, it will be obvious that it is necessary to establish two holdingpoints for traffic wishing to take-off from or to cross the ILS runway in use. In low visibilityconditions aircraft and other traffic will be required to hold well clear of the runway in use (at thecategory 2/3 holding point). In less critical weather conditions traffic may be permitted to hold at thecategory 1 holding point, which is closer to the ILS runway. The category 2/3 holding point may beused, at the discretion of ATC, when requested by pilots making an ILS approach in simulated lowvisibility conditions, despite the fact that the actual meteorological conditions do not in factnecessitate them.

The ILS Approach Plate73. The student should be familiar with the following definitions which are relevant to ILSprocedures.

Initial approach segment. That segment of an instrument approach procedure between theinitial approach fix and the intermediate approach fix or, where applicable, the final approach fix orpoint.

Intermediate approach segment. That segment of an instrument approach procedure betweeneither the intermediate approach fix and the final approach fix or point, or between the end of areversal, racetrack or dead reckoning track procedure and the final approach fix or point, asappropriate.

Final approach segment. That segment of an instrument approach procedure in whichalignment and descent for landing are accomplished.





Minimum sector altitude. The lowest altitude which may be used which will provide aminimum clearance of 300m (1000 ft) above all objects located in an area contained within a sectorof a circle of 46 km (25 nm) radius centred on a radio aid to navigation.

74. Figure 6-17 shows part of a typical ILS approach plate - in fact, the ILS approach ontorunway 24 at Manchester International. If applicable you will be fully briefed on the use of theapproach plates during your instrument rating flight training. For now, it will suffice to go brieflythrough the procedure as an initial familiarisation exercise.

75. Let us assume that the STAR (standard instrument arrival) procedure has been followed. TheILS procedure starts at 3000 ft (QNH) outbound on the reciprocal of the ILS QDM (237° - 180° =057°), descending to 2750 ft QNH once the aircraft is east of the outer marker/ME locator beacon.Having flown outbound from the outer marker for 30 seconds (corrected for wind, not specified onthe portion of the plate which is shown at Figure 6-17), the aircraft is then required to execute aprocedure turn (using an outbound track of 012°(M) to establish on the ILS QDM of 237°(M),maintaining an altitude of 2750 ft until intercepting the 3° glidepath at a MCT/DME range of 8 nm.

76. On crossing the OM inbound the height should be 1470 ft QNH, descending to DecisionHeight (DH). At DH the appropriate visual reference for landing must have been established and, ifit can be maintained, the aircraft may continue to descend. If not, then at DH the aircraft must climbahead on the missed approach procedure shown by dashed lines on the Plan and Elevation diagramat Figure 6-17.





FIGURE 6-17Typical ILS Approach Plate






Coverage

Data Communications (Special Information)

AZ Angular Measurement

EL Angular Measurement

Frequencies

Development of Multi-mode Receivers (MMR)

Positioning of Ground Transmitters

Sources of Error





7The Microwave Landing System

1. The present universal approach and landing aid is ILS. The requirement for a replacementsystem is justified when the shortfalls of ILS are realised, namely:

(a) ILS has a narrow, single approach path

(b) Only forty channels are available

(c) Signals, particularly those of the glideslope, are site and terrain sensitive

(d) With ILS it is only possible to radiate beams which define a single glideslope angle atany given installation. This makes the system inflexible in terms of, for example,helicopter and STOL aircraft operations.

2. In 1978 the Air Navigation Commission of ICAO evaluated four systems, as a replacementfor ILS. The system which was chosen was of American design and was known originally as theTime Referenced Scanning Beam (TRSB), but now more commonly as the Microwave LandingSystem (MLS). Until at least the early part of 2002 ILS will remain the ICAO non-visual standardaid.

3. MLS offers the following advantages over ILS :-

(a) not subject to the siting problems (buildings, terrain etc) that are inherent with ILS.

(b) elimination of ILS/FM broadcast interference problems;

(c) provision of all-weather coverage up to ± 60° from runway centreline, from 0.9° to15° in elevation, and out to 20 nautical miles (nm);





(d) capability to provide precision guidance to small landing areas such as roof-topheliports;

(e) continuous availability of a wide range of glidepaths to accommodate STOL andVTOL aircraft and helicopters;

(f) accommodation of both segmented and curved approaches;

(g) availability of 200 channels - five times more than ILS;

(h) potential reduction of Category I (CAT I) minimums;

(i) improved guidance quality with fewer flight path corrections required;

(j) provision of back-azimuth for missed approaches and departure guidance;

(k) elimination of service interruptions caused by snow accumulation; and

(l) lower site preparation, repair, and maintenance costs.

4. Like ILS, the system is based upon ground transmitters radiating information which isinterpreted by an aircraft receiver. The difference between ILS and MLS lies in the fact that, in MLSthe receiver calculates angles in both azimuth and in elevation by measuring the time intervalbetween successive passes of narrow (fan shaped) radiated beams. Ranging is derived from anaccurate DME installation, rendering marker beacons unnecessary.

5. The system may be divided into five functions :

(a) Approach azimuth;

(b) Back azimuth;

(c) Approach elevation;





(d) Range; and

(e) Data communications (Special Information)

6. With the exception of DME, all MLS signals are transmitted on a single frequency throughtime sharing. Two hundred channels are available between 5031 and 5090.7 MHz. By transmitting anarrow beam which sweeps across the coverage area at a fixed scan rate, both azimuth and elevationmay be calculated by an airborne receiver which measures the time interval between sweeps. For thepilot, the MLS presentation will be similar to ILS with the use of a standard CDI or multi-functiondisplay.

Coverage7. As with ILS, it is necessary to consider the coverage both in terms of azimuth and of elevationas well as the area of DME coverage.

Approach Azimuth Guidance8. The azimuth antenna provides lateral guidance during the approach. Azimuth station (AZ)coverage extends 40° on either side of the runway centreline, with a planned option of up to 60° oneither side of the centreline, and out to a range of 20 nm. The azimuth coverage is illustrated atFigure 7-1.





FIGURE 7-1Approach Azimuth Coverage





Back Azimuth Guidance9. The back azimuth antenna, if provided, gives lateral guidance for missed approach anddeparture navigation. The back azimuth transmitter is essentially the same as the approach azimuthtransmitter. However, the equipment transmits at a somewhat lower data rate because the guidanceaccuracy requirements are not as stringent as for the landing approach. The equipment operates onthe same frequency as the approach azimuth but at a different time in the transmission sequence. Onrunways that have MLS approaches at both ends, the azimuth equipment can be switched in theiroperation from the approach azimuth to the back azimuth and vice versa. Bi-directional MLSfacilities will have a separate DME/P and elevation TX (see later) for each direction of operation;only one DME/P and elevation TX will be operational at a time. If required, the left and rightcoverage can be asymmetric. The back azimuth coverage is illustrated at Figure 7-2.





FIGURE 7-2Back Azimuth Coverage





Elevation Guidance10. The elevation station (EL) transmits signals on the same frequency as the azimuth station. Theelevation station provides a wide range of glidepath angles. The glidepath angle which is requiredfor the approach by a specific aircraft is selected by the pilot. The EL signal coverage extendsthrough the AZ coverage, and so provides precision glidepath guidance at all points where azimuthguidance is available. Figure 7-3 below shows the EL coverage.





FIGURE 7-3Elevation Coverage





Range Guidance11. The precision DME (DME/P) provides continuous range information out to a range of 22 nmomnidirectionally. Accuracy is in the order of ±100 ft during final approach. The principle ofoperation is the same as conventional DME, but since no new frequency allocations are available, the200 channel capability has been provided by pairing with existing ILS installations and by theadoption of further values of pulse-pair spacing.

Data Communications (Special Information)12. The azimuth ground station includes data transmission in its signal format which includesboth basic (i.e. system data) and auxillary data (i.e. approach conditions). Basic data may includeapproach azimuth track and minimum glidepath angle. Auxillary data may include such informationas runway condition, windshear or weather.

AZ Angular Measurement13. The principle of azimuth angular measurement is shown at Figure 7-4.





14. The azimuth signal is a narrow vertical fan-shaped beam which sweeps back and forth acrossthe coverage area. Seen from the approach side, the beam starts at the left and sweeps at a uniformrate (constant angular velocity) to the right. This is known as the TO scan. After a short period,known as the guard time, the beam sweeps back to the starting point, and this is known as the FROscan (derived from ‘to and from’). Thus, within the complete cycle of the TO and FRO scan, twopulses will be received by the aircraft, and it is the accurately calculated time interval between thesetwo pulses which is proportional to the angular location (in azimuth) of the aircraft. The rate of scanis 13.5 scan cycles per second. Obviously, accurate time referencing is required and the TO scan ispreceded by what is known as preamble information which comprises basically a receiver referencetime code and a function identity code.

15. At Figure 7-4 an aircraft to the left of the centreline measures a certain time interval betweenthe TO and the FRO passage of the beam. At Figure 7-5, where the aircraft is to the right of thecentreline, the time interval between the passage of the TO beam and the FRO beam is reduced.

16. It can therefore be seen that the maximum time interval between the passage of the TO andFRO beams would be measured when the aircraft is at the extreme left edge of coverage andminimum time when the aircraft is at the extreme right hand edge of coverage. The measured timeinterval represents angular position and therefore displacement from the centreline. Thisinformation is supplied to a CDI or similar MLS display.





FIGURE 7-4The Principle of Azimuth Angular Measurement





FIGURE 7-5The Principle of Azimuth Angular Measurement

EL Angular Measurement17. The EL scanning principle is the same as that of the AZ. The beam is now a narrowhorizontal beam sweeping up and down at 40.5 scan cycles per second.





18. By combining the elevation information with the azimuth and range information previouslydiscussed, it is now possible to determine a 3 dimensional position which can be used by the on-board equipment to compute steering commands in relation to curved approaches, varyingglideslopes, segmented approaches etc. In the case of curved or segmented approaches the steeringinformation will be provided by two cross bars, directed by a computer which has been programmedwith the precise approach path to be flown.

19. It will be appreciated that in the absence of a DME/P signal only two dimensionalinformation will be available to the pilot, and it will therefore only be possible to fly straight-inapproaches in a similar manner to normal ILS approaches.

Frequencies20. The number of available channels is 200, spaced 300 KHz apart from 5031.0 MHz to 5090.7MHz (SHF). Since the basic technique is that of time multiplexing, (illustrated at Figure 7-6), allfunctions can take place on a single channel. Essentially, each function in the time-spacedtransmission format is a separate entity and is preceded by identification preamble. The receiver cantherefore recognise each element of the sequential transmission. The emission designators used forMLS are NOX (the unmodulated carrier) and G1D (the data transmission component).

21. As mentioned earlier Basic data words include the station indentification (four characterdesignator starting with the letter M) as well as the digital data needed by the receiver for processingthe azimuth, back-azimuth and elevation angle functions. Auxillary data words may contain suchinformation as system condition, runway condition and weather. Both Basic and Auxillary datawords are transmitted at a rate of once per second.





FIGURE 7-6The Multiplexing of MLS Information

Development of Multi-mode Receivers (MMR)22. The Microwave Landing System (MLS) and the Global Positioning System (GPS-see Chapter17) are both navigation programs that will play major roles in future Air Traffic Control Systems.For example, the FAA is committed to the implementation of MLS for precision approaches, and ispursuing an extensive program to determine the capabilities of GPS for use during all phases offlight.

23. GPS has a primary role to provide en-route navigation, however, in the terminal area GPS canprovide guidance to the MLS coverage area. GPS also has the capability to replace or supplementsome of the MLS functions, for example, it could provide the ranging information replacing DME/P.

24. Since ILS equipment is unlikely to be totally replaced by MLS for a number of years, multi-mode receivers have now been developed which have a fully integrated ILS/MLS/GPS capabilitytogether with Area Navigation (RNAV) facilities. (See Chapter 12).

Positioning of Ground Transmitters25. The diagram at Figure 7-7 shows the position of the MLS transmitters.





FIGURE 7-7Positioning of MLS Transmitters





Sources of Error26. Critical areas are regions around the MLS transmitters where vehicles and aircraft may causesignal errors as a result either of shadowing (where the offending aircraft or vehicle interrupts thesignal path) or multipath transmissions (where the signal is reflected by the offending aircraft orvehicle).

27. To minimise shadowing it is proposed to site the EL transmitter on the opposite side of therunway from the entry taxiway and the small transmitter size facilitates this. Multipath or reflectedsignals are more difficult to address. MLS uses a wavelength of 6 cm and therefore small flatsurfaces can produce high intensity reflections. Fortunately, these reflections tend to be highlyvariable in amplitude and in duration. The MLS receiver design is such that acquisition andvalidation circuits are able to select and process the strongest and most persistent signal, therebyminimising multipath errors.

NOTE:

Now read Chapter 10






Primary Radar

Secondary Radar

Continuous Wave Radar





8Basic Radar Principles

1. There are many applications for radar in Civil Aviation. A list of the most common uses isgiven at Figure 8-1.

FIGURE 8-1Ground Radar Types, Wavelengths and Frequencies

Type Frequency Wavelength

ATC Radars

Surveillance Radar 600/1300/3000 MHz 50/23/10 cm

Secondary Surveillance Radar 1090/1030 MHz 27 cms

Precision Approach Radar 9-10 GHz 3-3.3 cms

Airborne Radars

AWR/Mapping Radar 9.375 GHz 3.2 cm

Radio Altimeter 4.2 to=4.4 GHz 7 cm

Doppler 8.75 to 8.85 GHz13.25 to 13.4 GHz

3.4 - 2.25 cm

DME 960 to 1213 MHz 27 cms

Meteorological Radar

Weather Detection Radars 3000 MHz9-10 GHz

10 cm3-3.3 cm





2. There are basically two types of radar currently in use, pulsed systems and continuous wavesystems. Apart from a brief comparison of the properties of continuous wave radar and pulsed radarat the end of this chapter, the following paragraphs deal mainly with pulsed radar principles andpulsed radar systems.

3. The principle of pulse modulation was mentioned briefly in the Radio Theory part of thesyllabus where it was described as a short burst of electro-magnetic energy followed by a relativelylong quiescent period during which the transmitter is inactive.

4. A typical system might operate on a carrier wave frequency of 10 GHz with a pulse repetitionfrequency (PRF) of 1000 pulses per second (PPS) and a pulse width of 1 µ sec (one millionth of asecond or one microsecond).

5. Appreciate firstly that each pulse will contain 10,000 cycles of electro-magnetic energy (1 x

10-6 seconds at 10 x 109 Hz), and secondly that the relatively long period of transmitter quiescence isin this case only very slightly less than one thousandth of a second.

Primary Radar6. With primary radars a continuous train of pulses is beamed from the transmitter, via theaerial, into the atmosphere. In the event that these pulses strike a target, a small proportion of thetransmitted energy will hopefully be reflected back to the aerial and fed into the receiver. Thedirection in which the aerial is pointing at this time denotes the bearing of the target, whilst the timebetween transmission and reception of each individual pulse is used to determine target range fromthe radar head.





Beaming Technique7. If a primary radar system is to accurately determine target bearing, and to achieve optimumrange for a given transmitter power output, it is necessary that the beam of electro-magnetic pulsesbe as narrow as possible. The reader will undoubtedly have seen radar dishes of varying shapes andsizes, however most dish aerials are based on the geometry of the parabola. The essence of such anaerial arrangement is that the energy is transmitted from the antenna back into the dish. If theantenna is precisely positioned, all of the energy striking the dish will reflect into a parallel-sidedbeam. Unfortunately the antenna is necessarily of finite dimensions and so the beam is not quiteparallel-sided and spreads slightly giving the beam width shown at Figure 8-2.

FIGURE 8-2Typical Radiation Pattern





8. With pulsed systems it is normal to use one aerial system for both transmission and receptionof the pulses. It is necessary to disconnect the receiver from the aerial whilst the transmitter is firingand this renders primary pulsed systems unsuitable for very short-range systems, since the leadingedge of the pulse will arrive back at the radar head whilst the transmitter is still firing and will nottherefore reach the receiver. In other words, it is the pulse length which governs the minimum range.The longer the pulse length the greater the minimum range.

9. If the pulse length of a radar were 2µ sec that would mean that a received pulse could notenter the receiver for this amount of time after the start of the transmitted pulse. The minimum rangeof this radar would therefore be;

NOTE:

12.36 µ sec is the time taken for a pulse to travel out to a target 1 nm from theradar head, and back again-see paragraph 22.

10. In order to achieve a realistic range a narrow beam of pulsed energy is required and this maybe achieved by using the directly fed parabolic reflector described above, or a flat plate planar array.For a given diameter and wavelength the flat plate aerial provides a higher gain, a narrower beamand the least side lobe power.

212.36-------------nm approximately1000 ft=





11. The flat plate aerial produces a very narrow beam which is normal to the plane of the plate.With the parabolic reflector, there is a considerable amount of energy ‘spilled’ out of the aerial whichleads to the possibility of ground returns (see Figure 8-2). A further factor to consider is that thelarger the dish, the narrower the beam width, hence at a frequency of 9 GHz a 12 inch diameter dishwill give a 7° beamwidth, whereas a 30 inch diameter dish will give a 3° beamwidth.

12. The beamwidth of an aerial can be calculated relatively easily by using the following formula:

Transmitter Power vs Range13. The pulse from a primary radar system not only has to travel out to a target, but must travelan equal distance back to the receiver with enough strength to overcome any receiver noise. Theformula which relates power to range is:

14. In other words, in order to double the range of a radar the power must be increased by 16times its original value.

Beamwidth (in degrees) =

where

D

= wavelength in use

diameter of aerial dish Both in the same units=

70 λTX×D

----------------------

λTX

max range α power4





Range Determination15. The first thing to appreciate with regard to range determination is that, regardless of whetherthe radar head is ground-based and the target is airborne, or the radar head is airborne and the targeton the ground, it is SLANT RANGE which is determined. Figure 8-3 illustrates this fact, showing inthis case a ground-based radar head and an airborne target.

FIGURE 8-3Slant Range Determination in a Radar





16. Mathematically the determination of range is simple, since the speed of propagation is a

known constant (3 x 108 metres/second) and distance = speed x time. Remember however that thedistance involved is for the return journey, and so the target slant range is half of the distance in theabove formula.

17. The maximum range which can be achieved by any primary pulsed radar depends on severalfactors, one of which is the pulse repetition frequency employed. For example, ignoring pulse length,if the radar were firing 2 pulses per second then each pulse would have only half a second for the outand back journey before the next pulse was fired. More typically, if the radar were firing 1000 pulsesper second then each pulse would only have one thousandth of a second (1000 µ sec) for the out andback journey before the next pulse was fired. In other words, the pulse repetition frequency affectsthe maximum range of a pulsed radar, the higher PRF the shorter the maximum range.

18. Other factors which affect range, since radar equipments invariably operate at UHF or above,are:

(a) the height of the radar head

(b) the height of the target

(c) the presence of intervening high ground

and additionally such factors as:

(d) the power transmitted, and the beam width

(e) the nature of the target in terms of material, size, shape and aspect





Antenna Rotation Speed19. In those applications where a radar antenna is required to rotate through 360° in order tosearch for a target (e.g. a surveillance radar) there is a need to select the rotation speed very carefully.The optimum scan rate will be related to the following factors:

(a) pulse duration

(b) pulse repetition frequency (PRF)

(c) transmission power

20. The selection of antenna rotation speed is described in brief in Chapter 9.

Primary Radar Range Calculations21. To cope with these calculations, simply remember that:

(i) The distance travelled by the pulse (metres)

= Speed of propagation (metres/sec)

x

Time between transmission and reception of the pulse

(ii) The target range = Half the distance travelled by the pulse





(iii) The maximum range of a primary radar (ignoring pulse width and timebase flyback, to be covered later) is governed by the pulse repetition frequency (PRF). In order to avoid range ambiguity each pulse must return to the radar head before the following pulse is transmitted.

In other words the time taken for the return journey of the pulse must not exceed

seconds.

1PRF-----------





EXAMPLE 8-1EXAMPLE

The time between transmission and reception of a single pulse is 300 microseconds. Determinethe range of the target.

SOLUTION

Distance =

2 x range =

Range=

Range=

Range =

Range =

speed of propagation (m/sec) time (sec)×

speed of propagation (m/sec) time (sec)×

speed of propagation (m/sec) time (sec)×2

----------------------------------------------------------------------------------------------------

300,000,000 (m/sec) 300 µsec( )×2 1 000,000,×-----------------------------------------------------------------------------------

45,000 metres

45km





EXAMPLE 8-2EXAMPLE

Ignoring pulse width and flyback, calculate the maximum range in nautical miles for a primaryradar having a PRF of 500 pulses per second.

SOLUTION

2 x maximum range (metres) =

2 x maximum range=

maximum range=

=

=

=

speed of propagation (metres/sec) time (sec)×

speed of propagation1

PRF-----------×

speed of propagation2 PRF×

--------------------------------------------------

300 000,000,2 500×

-------------------------------

300 km

162 nm





EXAMPLE 8-3EXAMPLE

Ignoring pulse width and flyback, calculate the maximum permissible PRF for a primary radarwhich is required to give a range of 200 nm.

SOLUTION

Time=

=

PRF=

200nm = 370 km

PRF=

PRF=

PRF =

2 maximum range (metres)×speed of propagation (m/sec)----------------------------------------------------------------------

1PRF-----------

2 maximum range×speed of propagation---------------------------------------------------

speed of propagation2 maximum range×--------------------------------------------------

300 000 000, ,2 370 000,×---------------------------------

30 000,74

------------------

405 pulses per second





The Radar Mile22. The ‘radar mile’ will often be found to be a useful short cut to the basic calculation equationsshown in the previous examples. The radar mile is the time taken for a radar pulse to travel twonautical miles (that is to say one nautical mile out and one nautical mile back). Its value is 12.36 µ

sec (12.36 microseconds or 12.36 x 10-6 seconds). Thus, if the time between transmission andreception of a pulse were 100 µ sec, the target range (100 / 12.36) would be 8.1 nm. Appreciate thatthe answer is always in nautical miles (or alternatively the entry of distance must be in nauticalmiles).

Example 8-1 (reworked)

Range =

=

=

30012.36-------------

24.3 nm

45 km







Time =

=

Range =

=

Time =

=

PRF =

=

=

1PRF-----------

2000µ sec

200012.36-------------

162 nm

200 12.36×

2472µ sec

1time-----------

1,000,0002472

-------------------------

405pulses per second





Dead Time23. As mentioned earlier the choice of PRF determines the maximum range of a pulse radar. Inpractice the transmitted pulse does not stop at the maximum range and may therefore be reflectedback from a more distant target. A further period (known as the ‘dead time’) is therefore allowed forany echoes returning from targets beyond the specified range of the equipment.

24. Consider this example. A Terminal Area Surveillance Radar has a typical maximum range of75 nm. The time required for the radar pulse to complete this two way journey is 927 µ secs.

However, as the PRF of this type of radar is approximately 450 pps (2222 µ secs pulse spacing) this

allows for a dead time of 1295 µ secs. It is still possible, under certain meteorological conditions, toreceive intermittent responses during the dead time although, from a practical point of view, theycreate little real problem.

Primary Radar Displays25. Primary radar information is normally displayed on a cathode ray tube (CRT). Figure 8-4shows three aircraft in relation to a ground radar head the aircraft paints as they would appear on aplan position indicator (PPI) type of cathode tube display. Notice the distinctive tails of the targetpaints which give the radar operator some idea of target heading and speed.





FIGURE 8-4Plan Position Indicator (PPI) Display





26. The timebase on a PPI is synchronised with the aerial such that as the aerial sweeps throughnorth the timebase passes through north on the screen (normally screen vertical).

27. The timebase is produced by a beam of electrons striking the fluorescent coating on the faceof the screen. The timebase generator is synchronised with the transmitted pulse train such that asthe pulse leaves the transmitter the spot of light leaves the centre of the screen and travels radiallyoutwards at a linear rate, reaching the circumference of the screen a finite time later but before thenext pulse is fired. Because this spot of light is painting very rapidly on the screen the result appearsto the operator as an unbroken line rotating continuously about the screen at the same rate as theaerial is itself rotating.

28. As a return of energy is received via the target it is amplified and fed to the CRT causing amomentary increase in the electron flow producing the timebase, and therefore an increase in theintensity of the light spot. This brighter spot, which represents the target, will linger on the screenonce the timebase has passed, fading only slowly and being re-illuminated next time around. Sincethe rate at which the timebase is manufactured is linear, the distance of the target paint from thecentre of the screen accurately represents the actual (slant) range of the target from the radar head.Range rings may be etched onto the face of the CRT or electronically painted onto the screen. Pleasenote that on some radar systems it is possible to determine the position of an aircraft by readingbearing and distance off the radar screen with the aid of electronic devices such as electronic bearinglines and variable range rings.

The Radar Resolution Rectangle29. We have already discussed the desirability of using a narrow beam for pulsed radar systems,in order to achieve maximum range by greater concentration of power, and to accurately determinethe target bearing.





30. Similarly, we have already mentioned that the pulse width (or pulse length, the terms aresynonymous) will determine the minimum range of the system, whilst the pulse repetition frequencywill govern the maximum theoretical range.

31. It is now necessary to look briefly at how pulse width and beam width will distort the targetpaint on the screen of a plan position indicator type of display. The distortion which occurs becauseof these two factors (plus in fact one other consideration), is termed the radar resolution rectangle.

32. The PPI at Figure 8-4 shows a timebase originating at the centre of the screen and rotatingclockwise around the screen, in synchronisation with the transmitter aerial. Put another way, thetimebase is aligned with the centre of the beam. Unfortunately the beam has a finite width (whichincreases with range), and the target will start to paint on the screen when the leading (right hand)edge of the beam first illuminates the target, and will continue to paint until the rear (left hand) edgeof the beam finally ceases to illuminate the target. The lateral dimensions of the target will thereforebe distorted (stretched) by one whole beamwidth.

33. As an example of this particular problem consider a radar with a beamwidth of 4°. Using the1 in 60 rule it can be calculated that two targets at a range of 50 nm will be stretched by about 3½

nm, and would therefore also have to be separated by 3½ nm in order to show as separate returns.As the target range increases, the amount of spacing required between the aircraft will also need toincrease. The ability of a radar to discriminate in a lateral direction between two targets is calledazimuth resolution.





34. Similarly, the leading edge of the pulse, reflecting from the target and arriving back at thereceiver will determine the range at which the target first starts to paint on the screen. The target willcontinue to paint on the screen until the trailing edge of the pulse has ceased to illuminate the target.This will result in a range (or depth) distortion of the target of a distance equivalent to half the pulse

width, converted into metres, using a speed of 300 metres per micro-second (3 x 108metres / second).The depth distortion is governed by half the pulse width, since the time involved covers the returnjourney.

35. If the pulse width/length in a primary radar was 2 µ sec, the target depth (range) distortioncould be calculated as follows:

36. To put it another way, if two targets are on the same bearing but within 984 ft of each otherthey will not show as separate returns. The ability of a radar to discriminate in a ranging sensebetween two targets is called radial resolution.

37. The target will be further enlarged, both in width and depth, by the radius of the spot of lightwhich is used to generate the visual timebase, and to paint the target.

38. Modern air traffic control PPI radars overcome the problems of target distortion due to theresolution rectangle, and the consequent tendency for two adjacent targets to overlap and paint asone on the screen, by suppressing the target paints entirely and replacing them with electronicallyproduced crosses. Additionally you would expect to see the information supplied by the SSRequipment and a map of the airways structure on such a sophisticated display.

Depth Distortion2µsec

12.36µsec-------------------------�� nm 984ft= =





Other Factors Affecting Quality of Target Depiction on a PPI Display39. The following additional factors affect the quality of the target as displayed on a PPI display.

Super Refraction. Under certain meteorological conditions the detection range of objects close tothe Earth’s surface can be considerably increased. This process is called super refraction and is morefully described in Chapter 9.

Sub Refraction. Again, under certain meteorological conditions the detection range of objectsclose to the Earth’s surface can be considerably reduced. This process is called sub refraction and ismore fully described in Chapter 9.

Attenuation with Distance. As a radar pulse travels out from the radar head its strength willweaken due to atmospheric attenuation. Consequently the greater a given target’s range, the smallerwill be the amount of reflected energy that will be returned to the antenna.

Condition and Size of Reflecting Surface. Factors that determine the amount of reflectedenergy from a given target are the size and shape of the reflecting surface, the actual material that thereflecting surface is made from (e.g. metal reflects better than wood), and the aspect of the target (e.g.an aircraft flying directly towards a radar head will reflect less energy than an aircraft flying at aconstant range). If an aircraft changes attitude whilst being illuminated by a radar there may be achange of polarisation of the radio wave, which could result in the target fading from the radarscreen (the stealth bomber utilises this fact to become ‘invisible’ to a radar system).





Moving Target Indication (MTI) A technique called Moving Target Indication may be used toreduce clutter on a PPI screen generated by echoes from permanent objects such as hills, buildingsetc. MTI is more fully discussed in Chapter 9.

Secondary Radar40. Secondary radar does not rely on reflections of the interrogation pulse arriving back at theradar head via the target. Instead, a booster transmitter or transponder is situated at the target andthis is used to revitalise the interrogation pulse for the return journey. Obviously such a systemrequires the co-operation of the target. Two examples of secondary radar will be considered shortly,namely SSR and DME.

41. Secondary radar has the following advantages when compared with primary radar:

(a) Since only enough energy need be transmitted for a one-way journey, the requiredtransmitter power is lower and consequently the equipment lighter and less bulky.

(b) Pulse sequences may be coded, thereby conveying additional intelligence, for examplemode C (pressure altitude readout) with SSR.

Continuous Wave Radar42. The only continuous wave primary radio/radar system presently considered in this syllabus isthe radio altimeter.





43. Pulsed radar systems use a single antenna which is switched in turn between transmitter andreceiver. It therefore follows that transmission and reception cannot take place at the same time in apulsed radar system. This leads to the minimum range problems of pulsed systems, especially thoseusing wide pulse widths. There is a dark area around the radar antenna and aircraft flying in thisarea do not paint on the display, since the returns from these targets arrive back at the antenna whilstit is still switched to the transmitter.

44. Continuous wave radars transmit and receive continuously and therefore have separateantennae for each function. Consequently the receiver is always on line and therefore no minimumrange problem exists.

45. Pulsed radars determine the range of a target by measuring the time taken for a transmitterpulse to travel to the target and to return. Since continuous wave radars, by definition, do not usepulses, it would appear that range determination is impossible. In fact, very accurate rangedetermination is achieved (for example in the radio altimeter) by frequency modulating thecontinuously transmitted signal and by then comparing the frequencies of the transmitted and thereceived signals at precisely the same point in time.

46. Continuous wave radars transmit much lower power signals than pulsed radars. A furtheradvantage of continuous wave systems is that they can operate with a much narrower bandwidth,consequently a better signal to noise ratio is more easily attainable.





Ground Based Radars

En-Route Surveillance Radar (RSR)

Airfield Radars

Airport Surveillance Detection Equipment (ASDE)

Second Trace Returns

Accuracy of Ground Based Radars




Ground Based Radars


9Ground Based Radars

1. Ground Based Radars can be generally divided into two groups: Long range radars or shortrange radars.

Long Range Radars employ lower frequencies (i.e. longer wavelengths, typically 10-20 cm),lower PRFs and larger pulses to give greater ranges with less attenuation. Antenna rotation rates arelow (5-15 rpm) as target movement at long ranges is relatively slow.

Short Range Radars use high frequencies (i.e. short wavelengths, typically 3 cm) to give shortrectangular pulses at relatively high PRFs for low minimum range, better resolution and greateraccuracy. Antenna rotation rates are high (up to 60 rpm) as target movement at short ranges isrelatively fast and frequent radar updates of position are therefore required.

2. Two categories of ground based radars are briefly discussed below, en-route surveillanceradars which are used for middle and upper airspace control, and airfield radars.

En-Route Surveillance Radar (RSR)3. En-route surveillance radars are used to monitor airways traffic at ranges up to 250 nm.Range and bearing information is provided by a Primary Radar, with a Secondary Surveillance Radar(SSR - refer to Chapter 11) providing additional information.

4. The preferred frequency for these radars is 600 MHz, giving a wavelength of 50 cm. At theserelatively long wavelengths, rain and weather present far less of a problem than with higherfrequency systems.



Ground Based Radars


Airfield Radars5. Larger airfields are normally equipped with ground-based primary radar systems. These willnormally comprise of a Terminal Area Surveillance Radar, with perhaps an additional PrecisionRadar system. Surveillance radars are installed at most major civil airports whereas Precision Radarsare more usually confined to military airfields, particularly in the UK.

Terminal Area Surveillance Radar (TAR)6. The surveillance radar consists of a scanner which rotates through 360 degrees in thehorizontal plane at between 5 and 15 rpm. Presentation of the radar picture is achieved by using aplan position indicator (PPI) which enables the controller to determine the aircraft's range andbearing from the airfield (but not its height). Once again this radar is normally supplemented by SSR.

7. Surveillance radars normally use a beam which has a horizontal width of one degree and avertical depth of 40 degrees. Three frequencies are commonly used:

(a) 3000 MHz - giving a wavelength of 10 cm.

(b) 1300 MHz - giving a wavelength of 23 cm.

(c) 600 MHz - giving a wavelength of 50 cm.

8. The higher the frequency, the smaller the aerial array for the desired beam width.Unfortunately, however, the higher the frequency the greater the likelihood of the screen becomingcluttered by weather returns. This weather clutter can be electronically suppressed with modernradars, but not without the loss of some degree of picture definition.



Ground Based Radars


9. The range of surveillance radars depends on the power transmitted, and of course on line ofsight considerations. Typically such systems achieve ranges of 75 nm.

Moving Target Indication (MTI)10. Some surveillance radars make use of the Doppler principle to eliminate radar returns fromfixed objects such as hills, buildings, masts and so on. This process is known as Moving TargetIndication (MTI). The principle is that returns from moving targets suffer a doppler shift, whereasthose returns from stationary targets do not - the radar only displays those that experience theDoppler shift. Since the Doppler principle requires that there be relative motion between a target anda transmitter, it will be obvious that a dangerous situation could occur when the target is maintaininga constant range from the radar head and therefore would not paint on the controller’s screen.

Surveillance Radar Approach (SRA)11. The surveillance radar may be utilised to give the pilot guidance during a descent to land inpoor visibility. During a surveillance radar approach (SRA) the PPI display provides the talk-downcontroller with the aircraft's range and bearing. By electronically or physically superimposing theextended centreline of the runway in use on the cathode ray tube, the controller is able to give thepilot fly left or fly right instructions in order to maintain the centreline. These instructions arequantified, for example:

(i) you are left of the centreline, turn right five degrees on to 262 degrees

(ii) maintain 262 degrees, closing centreline left to right

(iii) on the centreline, turn left three degrees on to 259 degrees

and so on.



Ground Based Radars


12. The surveillance radar gives no height information and consequently the descent is monitoredby the pilot with the aid of the controller who passes ranges from touchdown and check heights for,typically, a three degree glidepath. Surveillance radar approaches normally terminate at 2 nm fromtouch down, however when a high resolution radar is used, the surveillance radar approach mayterminate as close as ½ nm from touch down.

Precision Approach Radar (PAR)13. Airfields equipped with a precision radar system can offer a far more accurate talk-down,principally because the precision talk-down controller can monitor the height of the aircraft duringthe approach.

14. Precision radars normally operate in the 9 to 10 GHz frequency band, giving a 3.3 to 3 cmwavelength. These shorter wavelengths give the high definition required for precision approaches,but at these frequencies weather clutter presents a significant problem.

15. The precision approach system employs two independent radars and associated aerial systemswhich may be equated to the localiser and glidepath of the ILS.

16. The aircraft's approach is monitored in azimuth by a radar using a beam which is typicallyhalf a degree wide and two degrees in the vertical, and which scans 10 degrees either side of theextended runway centreline out to a distance of at least 9 nm. The radar return from the aircraft isdisplayed on a rectangular screen which has electronically superimposed upon it the extendedcentreline and range markers from the touchdown point.



Ground Based Radars


17. The aircraft's approach is monitored in elevation by a radar using a beam which is typicallytwo degrees in the horizontal and half a degree in the vertical, and which scans through seven degreesin the vertical plane from one degree below to six degrees above the horizontal plane. The radarreturn from the aircraft is displayed on a second rectangular screen which has electronicallysuperimposed upon it the glidepath (typically three degrees) and again range from touchdownmarkers, at one nautical mile intervals.

18. A typical PAR monitoring arrangement is illustrated at Figure 9-1 showing both the elevationand azimuth screens as they would be situated in front of the controller.



Ground Based Radars


FIGURE 9-1Typical PAR Monitoring System



Ground Based Radars


19. During a PAR (sometimes called a GCA-Ground Controlled Approach) the talk-downcontroller will give the pilot instructions to fly left or right together with instructions regarding hisrate of descent in relation to his position above or below the glidepath. The instructions in azimuthare again quantified but (in this country at least) the instructions in elevation are not, unless thecontroller wishes to emphasise to the pilot that he is dangerously low on the glidepath.

20. Since the aircraft is positively monitored in both azimuth and elevation during a PARapproach, the obstacle clearance height (OCH) and the pilot's decision height will be lower for aPAR than for an SRA.

Approach Radar Procedures21. The pilot of an aircraft requiring a radar approach at an airfield, or indeed radar vectoringfor an ILS or visual approach, should contact the airfield on the published approach frequency atleast 10 minutes before his ETA at the airfield. If it is the pilot's intention to fly a PAR or SRAapproach, the controller will hand the aircraft over to the director whose first job is to positivelyidentify the aircraft in question on the surveillance radar screen. This identification may be achievedin a number of ways:

(a) By using Secondary Surveillance Radar (SSR).

(b) By the director instructing the pilot to make identification turns and observing theseturns on the radar screen.

(c) By a pilot report over a designated reporting point.

(d) By the pilot obtaining a VOR/DME fix and passing this information to the director.

(e) By radar handover from another unit, or another controller within the same unit.



Ground Based Radars


Having identified the aircraft, the director will:

(a) Vector the aircraft, effecting separation from other aircraft, until the aircraft ispositioned for the PAR or SRA approach.

(b) Pass heights/altitudes to fly.

(c) Pass aerodrome information, including the weather.

(d) Pass the obstacle clearance height (OCH) for the type of approach to be used, and askthe pilot to check his decision height.

(e) Pass the radio failure procedure, if this is not published.

22. A typical radar circuit followed by a PAR approach is shown at Figure 9-2. Notice that thedirector normally hands over control of the aircraft to the precision talk-down controller atapproximately seven nautical miles from touchdown. By this time, the director will have positionedthe aircraft on the extended runway centreline.



Ground Based Radars


FIGURE 9-2Typical PAR Profile



Ground Based Radars


ILS Approaches Monitored by PAR23. Providing that PAR is available for the ILS runway in use, ILS approaches will be monitoredby PAR whenever the weather is below prescribed minima, or when requested by the pilot.

24. When monitoring an ILS approach, the precision controller will take no action unless:

(a) The aircraft strays outside the approach funnel which extends half a degree above andbelow the glidepath and two degrees either side of the centreline.

(b) A dangerous situation is seen to be developing.

(c) It appears certain that overshoot action may result if a certain action is not taken.

25. The PAR controller will at all times be prepared to convert the approach to a PAR talk-downif so requested by the pilot. The controller will terminate his monitoring of the approach when theaircraft is known to have landed or have gone around.

Airport Surveillance Detection Equipment (ASDE)26. Ground movement radars, which are frequently referred to as Airport Surveillance DetectionEquipment (ASDE) are installed at major aerodromes to control the safe movement of aircraft on theground, principally during low visibility operations (LVO).



Ground Based Radars


27. The latest generation of ground movement radars operate in the SHF band. This frequencygives the required target definition with acceptably low levels of weather clutter at the very shortranges required of the system. Earlier systems tended to operate in the EHF band which gaveexcellent target definition but which suffered unacceptable signal attenuation in precipitation. Thebeam width of the radar is narrow and the pulse length short, in order to minimise distortion of thetargets (see The Radar Resolution Rectangle in Chapter 8). The very high PRF employed alsoenhances target definition, and in any event is necessary due to the very high sweep rate which isemployed (60 rpm). This high sweep rate is used, not only to improve target definition, but also toshow the speed of movement of the target.

28. A comparison of the operating dynamics of two ground movement radars (Racal and Astre),together with typical equivalent values for ATC surveillance radars (dealing with airborne targets) isshown in the table at Figure 9-3.

FIGURE 9-3Parameter Comparison between GMRs and RSR/TAR

Type Wave-length Power Pulse length Scan rate Max range Beam width PRF

Racal 3 cm (SHF) 20 Kw 0.04 µ sec 60 rpm 2.5 nm 0.4° 4000 pps

Astre 1.8cm SHF 20 Kw 0.04 µ sec 60 rpm 6.5 nm 0.35° 8200 pps

RSR/TAR

50 cm to 10 cm 60 to 600Kw

2 to 5 µ sec 5 to 15 rpm

75 to 250nm

1° to 2° 250 to 1000 pps



Ground Based Radars


29. Ground movement radars are capable of differentiating between large, medium and smallaircraft, but not between specific types of aircraft. Large aircraft appear on the screen as distinctaircraft shapes, except that the wing which is shielded by the fuselage together with those parts of thewing which overhang the grass areas adjacent to taxiways, will not paint. Medium size aeroplanespaint as aircraft shaped returns in areas of good coverage, but in rough cruciform shape in otherareas. Small aircraft tend to paint as blips and could be confused with vehicles.

Second Trace Returns30. The choice of PRF in a ground radar is a compromise between two factors; too low a PRFwill mean that information updates on the radar screen are too slow, whilst too high a PRF gives aproblem due to second trace returns.

31. Suppose that having considered a radar’s parameters it can be calculated that its maximumdetection range is 200 nm. The total round trip time for a pulse travelling to a target at maximumrange would be approximately 2500 µ sec. If a PRF of 500 pps had been chosen for this radar (i.e.time interval between pulses of 2000 µ sec), the time base on the controller’s screen would start toregenerate 500 µ sec before the returning echo from the previous pulse, for a target at max range.This second trace return would therefore appear to be at 40 nms range (Figure 9-4).



Ground Based Radars


FIGURE 9-4Second Trace Returns

32. In the above example the maximum PRF that could have been used was 400 pps and in oldradars this was indeed a commonly used PRF. As the maximum range of radars has improved inrecent years it is now common place to find PRFs in the range 100 to 250 pps.

Accuracy of Ground Based Radars33. As described in the previous chapter, all of the above mentioned ground radar equipmentswill experience accuracy degradation, to a certain extent, due to azimuth and radial resolutionproblems. To be specific, the target size in the azimuth direction (i.e. lateral dimension) will beincreased by one whole beamwidth (the amount of distortion therefore increases with range from theradar head). The target size in the radial direction (i.e. range) will be increased by half the pulse

width, converted to a distance using .3 108× m/sec



Ground Based Radars


Factors Affecting Range and Accuracy34. The following factors affect the range and accuracy of ground radars:

Super-Refraction. Under certain meteorological conditions, radio waves in the VHF, UHF andSHF bands, which normally travel only in straight lines, may behave in a way which is at first sightsimilar to skywaves.

The meteorological conditions required for this type of propogation (duct propogation) are a markedtemperature inversion and a rapid decrease in humidity with height. Figure 9-5 shows ducting which,in this case, is occurring between the surface and a low level inversion. The signal is effectivelytrapped under the inversion and may travel hundreds of miles with little attenuation. In this way,when high pressure systems prevail, signals may be received from distant SHF transmitters which arefar beyond the normal direct wave range.

FIGURE 9-5The Ducted Wave (Super Refraction)

The process of ‘ducting’ is also known as super refraction and as such it can extend the detectionrange of a ground based radar system.



Ground Based Radars


Sub-Refraction. Sub-refraction is a condition of atmospheric refraction, created by gradients oftemperature and humidity, when radio waves are bent less than normal. This reduction in bendingwill result in a much smaller horizon distance than would be the case under normal propogationconditions.

Absorption and Reflection by precipitation. Precipitation is a cause of both absorption andreflection of radio energy. In general terms, as frequency increases (and wavelength decreases) theamount of absorption by precipitation increases, whereas an increase in frequency will cause theamount of reflected energy to increase as well. It is for this reason that airborne weather radarsoperate with compromise wavelengths between 3-10 cm; at wavelengths below 3 cm there is toomuch absorption of the radio wave, on the other hand at wavelengths above 10 cm there is too littlereflected energy. As a general statement it can be said that a radar pulse will reflect most energy fromwater droplets whose size is compatible to the wavelength in use.



Ground Based Radars




The maximum range of a pulsed radar is governed by which parameter?

QUESTION 2.

What is the formula for calculating the beamwidth of an aerial?

QUESTION 3.

Ignoring flyback, the maximum theoretical range of a primary radar with a prf of 800 pps is:

QUESTION 4.

A radar with a maximum range of 265 nm will have a maximum theoretical prf of:

QUESTION 5.

Given a prf of 3200 pps, the maximum theoretical radar range is:

QUESTION 6.

A primary radar system is required to have a maximum range of 400 nm. Ignoring pulse width andflyback, the maximum prf that the system could employ is:

QUESTION 7.

The advantage of CW radar over a pulse system is that:



Ground Based Radars


QUESTION 8.

A pulse radar gives best target definition with:

QUESTION 9.

What is the purpose of having 'Dead Time' in a radar?

QUESTION 10.

What are the two items of information that can be determined about a target from a simple primaryradar PPI display?

QUESTION 11.

If the pulse length in a particular primary radar was 4 microseconds, what would be the amount oftarget depth distortion?

QUESTION 12.

What principle does the Moving Target Indication (MTI) facility (as used in a ground radar) utilise?

QUESTION 13.

What are two the advantages of secondary radar when compared to primary radar?

QUESTION 14.

What frequency do En-Route Surveillance Radars (RSR) operate at and why?



Ground Based Radars


QUESTION 15.

At what range does a Surveillance Radar Approach (SRA) normally terminate?

QUESTION 16.

What frequency does PAR normally operate at, and what is the approximate wavelength?

QUESTION 17.

In a PAR what is the area of sweep of the azimuth radar?

QUESTION 18.

What guidance information is passed by the talk-down controller to an aircraft flying a PARletdown?

QUESTION 19.

What is the typical scan rate of an RSR/TAR?

QUESTION 20.

What is the name given to the phenomenon where radio waves are bent less than normal due totemperature and humidity gradients?



Ground Based Radars


ANSWERS:ANSWER 1.

Pulse repetition frequency (PRF)

ANSWER 2.

Beamwidth (in degrees) =

where:

= wavelength in use

D = diameter of aerial dish (with both in the same units)

ANSWER 3.

101 nm

ANSWER 4.

305 pps

ANSWER 5.

25 nm

ANSWER 6.

203 pps

70 λTx×D

----------------------

λTx



Ground Based Radars


ANSWER 7.

There is no minimum range problem

ANSWER 8.

Short pulse/narrow beam

ANSWER 9.

Dead time is a portion of the period of time between pulses which allows for any echoes returningfrom targets beyond the specified range of the equipment

ANSWER 10.

Slant range and bearing.

ANSWER 11.

Depth Distortion =

= 1968 ft

ANSWER 12.

The fact that there will be a Doppler shift in frequency in the reflected energy from a moving target.

4µ sec12.36µ sec--------------------------�� nm



Ground Based Radars


ANSWER 13.

Lower power requirements and the option of conveying additional intelligence by means of pulsecoding.

ANSWER 14.

600 Mhz - because rain and weather are less of a problem at this frequency.

ANSWER 15.

2 NM

ANSWER 16.

9-10 Ghz, wavelength approximately 3cm.

ANSWER 17.

10 degrees either side of the extended runway centreline.

ANSWER 18.

Fly left/right instructions together with instructions regarding rate of descent in relation to theaircraft position above or below the glidepath.

ANSWER 19.

5-15 rpm.



Ground Based Radars


ANSWER 20.

Sub-Refraction.






Weather Radar Frequency

The Aerial

The Control Unit

Hill Shadow

Functional Check of AWR

Determining the Height of Cloud Tops

Using a Monochrome AWR for Weather Avoidance

Use of AWR for Navigation Position Fixing

Coloured Screen Weather Radars

Factors Affecting the Range of an AWR





10Airborne Weather Radar

1. Airborne weather radar is a primary radar system, which as the name suggests is designed todetermine the presence ahead of the aircraft of hazardous weather, namely turbulent cloud. Theradar may also be used to paint a radar map of the ground features ahead of the aircraft.

Weather Radar Frequency2. The most commonly used transmitter frequency for airborne weather radar systems is 9375MHz in the SHF band. This frequency gives a wavelength of just over three centimetres.

3. In order to detect turbulent cloud the radar must receive reasonably strong target returnsfrom the large water droplets contained in the strong upcurrents associated with the turbulence.

4. The ability of water droplets to act as efficient targets depends upon their size relative towavelength of the transmitted frequency. The larger the target droplet in relation to the transmittedsignal wavelength, the better the return. A 3 cm wavelength gives a good return from these largewater droplets, but no significant return from the smaller water droplets associated with non-turbulent cloud. If a higher frequency were to be used (with a correspondingly shorter wavelength),the signal would be scattered by small water droplets, reducing effective range and cluttering thescreen with unwanted returns. Conversely, a lower frequency would be of no use since even largewater droplets would not give a satisfactory paint on the screen. Note that AWR will not detect clearair turbulence (CAT) since it is reliant on the presence of water droplets for its operation.





The Aerial5. In order to achieve a realistic range a narrow beam of pulsed energy is required and this maybe achieved by using a either a directly fed parabolic reflector, or a flat plate planar array. For a givendiameter and wavelength the flat plate aerial provides a higher gain, a narrower beam and the leastside lobe power. Since the flat plate array is approximately twice as efficient as the parabolic reflectorit is invariably used in a modern AWR system.

6. The flat plate aerial described above produces a very narrow beam which is normal to theplane of the plate. Note that when using the cheaper parabolic reflector the energy does not radiatefrom a point source at the focal point of the aerial, but generally from a dipole feed which gives anarrow, slightly diverging beam, which is circular in cross-section, and which is often referred to aseither a conical or pencil beam. With the parabolic reflector, there is a considerable amount of energy‘spilled’ out of the aerial which leads to the possibility of ground returns (see paragraph 8). A furtherfactor to consider is that the larger the dish, the narrower the beam width, hence a 12 inch diameterdish will give a 7° beamwidth, whereas a 30 inch diameter dish will give a 3° beamwidth at the sametransmitted frequency (9375 MHz).

7. The beamwidth of an AWR aerial can be calculated relatively easily by using the followingformula:

=

=

=

Beamwidth (in degrees) 70 λTX×D

----------------------

whereλTX wavelength in useboth in the same units

D diameter of aerial dish





8. The side lobe of energy escaping downwards from the dish serves a useful purpose on thisequipment. If the radar is searching ahead for turbulent cloud but none is present, the screen will beclear of target echoes. It may appear that the equipment is unserviceable, however this is easilychecked because, if the equipment is functioning correctly, the downward side lobe will cause aheight ring to be painted on the screen at a range corresponding to the height above the ground of theaircraft (except, perhaps, over a smooth water surface).

9. The scanner on an airborne weather radar is located in the nose of the aircraft and scanstypically from 45 to 60 degrees on either side of the aircraft centre-line. The timebase on the weatherradar screen is synchronised with the aerial.

10. There are two fundamental types of airborne weather radar currently in use. The earlier typeuses a monochrome screen, usually giving a green or an amber paint. Later radars present a colouredpaint and use the ascending colours of green, yellow and red to distinguish between light, mediumand heavy target returns (i.e. increasing intensity of precipitation). We will initially consider themonochrome option and subsequently discuss colour weather radars.

The Control Unit11. An AWR control panel is illustrated at Figure 10-1. The purpose served by the variouscontrols is discussed below.





FIGURE 10-1Typical AWR Control Panel

Power Switch12. This control is used in conjunction with the timebase range switch, especially during theequipment switching-on procedure. With the POWER switch ON (either on STAB ON or STABOFF) and the RANGE switch in the STANDBY position, the equipment is brought up to operatingtemperature but does not transmit.





Timebase Range Switch13. With the POWER switch ON the RANGE switch may be selected to the required range onceairborne and clear of the ground, the transmitter will fire immediately. The range options aretypically 20 nm, 50 nm and 150 nm.

14. Figure 10-2 shows typical range markers for these options. The range markers areelectronically superimposed on the screen and their brilliance may be adjusted using the MARKERBRILLIANCE control. The radial lines illustrated are etched on to the surface of the screen. Theselines represent angular deviation from the zero degrees relative position which is of course theaircraft extended centreline.





FIGURE 10-2Typical AWR Range Options





Tilt Control15. The scanner sweeps from side to side in a plane of elevation which is selected by the operatorusing the tilt control. This control can tilt the scanner through 30 degrees vertically, 15 degreeseither side of level. The level datum will either be the aircraft's yawing plane with the POWERswitch in the POWER ON - STAB. OFF position, or Earth horizontal with the POWER switch inthe POWER ON - STAB. ON position, see Figure 10-3. Note that the scanner is stabilised in bothpitch and roll.





FIGURE 10-3Effect of ‘Stab on/stab off ’ Selection





Function Switch16. The FUNCTION switch has four positions. The WEA (weather) and CONT (contour)modes are designed for cloud detection, whilst the MAN (manual) and MAP (mapping) modes aredesigned for ground mapping. These functions are discussed below.

WEA In the WEATHER mode the narrow conical beam is used. The manual gain control isinoperative and an automatic gain control circuit operates. The equipment is designed to paint onlyturbulent cloud on the screen. At close range even non-turbulent cloud may reflect enough energy tocause a paint. The sudden appearance of apparently turbulent cloud at close range might alarm andconfuse the pilot. To avoid this, the automatic gain control circuitry reduces the sensitivity of thereceiver progressively from 20 nm range down to the minimum range of the equipment. Rememberthat beyond this range the strength of return will become weaker as range increases.

CONT The type of AWR considered here is fitted with a monochrome screen. On this type ofequipment the difference between very turbulent cloud and less turbulent cloud is simply a differencein the intensity of the target paint on the screen. Even with a modern daylight screen, distinguishingbetween intense and less intense paints is almost impossible for the human eye. It is necessary toincorporate a CONTOUR mode.

The function of the iso-echo circuitry, which functions when the contour mode is selected, is to invertthe target signal above a given level. This results in particularly turbulent cloud painting on thescreen with a hole in the middle, the hole indicating the area of intense turbulence.

Figure 10-4 shows a high amplitude signal (turbulent cloud) and the resulting display with theequipment in the WEA mode. Figure 10-5 shows the signal amplitude and resulting display for thesame cloud, but with the equipment operating in the CONT mode.





FIGURE 10-4Portrayal of Information in ‘WEA’ Mode





FIGURE 10-5Portrayal of Information in ‘Cont’ Mode

MAN In the MANUAL mode the radar is used for long-range ground mapping, typically atranges in excess of 60 to 70 nm. In this mode the narrow conical beam is used to achieve therequired range but now the automatic gain control is inoperative and the MAN GAIN (manual gain)control is used by the operator to achieve the best picture definition.





MAP For short-range mapping the MAP mode should be selected on the function switch. Spoilersare now introduced into the dish of the aerial which distort the beam into the fan shape illustrated atFigure 10-6. Now a much greater area of the ground ahead is covered by the beam but of course theavailable power is spread over a correspondingly greater area thereby reducing the effective range.Again the automatic gain control circuits are inoperative and the manual gain control is used to besteffect. Because the strength of the returns at say 50 nm range, would normally be weaker thanreturns from the same sized object at 20 nm range, (due to the extra distance travelled by thetransmitted and reflected energy), the power distribution throughout this beam is varied so that avalid comparison of targets can be made by the operator. The power spread is adjusted so thatmaximum power is directed to the front of the beam, and thereafter is progressively reduced asdistances decrease, so that the power directed to the closest object is minimum. The reduction inpower with decreasing range is a function of the cosecant of the depression angle and the beam soproduced is sometimes referred to as a cosecant² beam. This beam shape is best for mapping sinceit enables returns over a wide area to be displayed, allowing the necessary cross checking between thechart and the display to identify fixing points. However, because of the power spread, returnsbeyond about 70 nm tend to be weak and it is then preferable to use the pencil beam for positionfixing.





FIGURE 10-6Beam Shape in Map Mode (Fan or

Cosecant2 Beam)

17. It will probably be fairly obvious that, when using the ground mapping modes of a AWR toilluminate a particular ground feature, the tilt setting must be increased in the downwards directionas the selected range decreases. Similarly the tilt setting will also have to be increased in thedownwards direction if the aircraft climbs to a higher altitude.

Hill Shadow18. Figure 10-7 shows a mapping beam being used over mountainous terrain. The ground in theshadow of the closest mountain is not being swept by the beam and will not therefore paint on thescreen. This could be interpreted erroneously as an indication of the presence of a stretch of waterwhere none would be expected. This effect is known as hill shadow.





FIGURE 10-7The Problem of Hill Shadow in an AWR

Functional Check of AWR19. It is obviously advisable to check the serviceability of the AWR on the ground prior to flight.When conducting these checks, certain precautions must be observed to avoid damage to personnel,ground installations and to the AWR itself. The following precautions are worthy of note:

(a) Ensure that the POWER switch is OFF prior to engine start to avoid surge currentswhich may damage the equipment.

(b) Following engine start, ensure that the RANGE switch is on STANDBY beforeturning the POWER switch to the ON - STAB OFF position. This will allow theequipment to reach its normal operating temperature.





(c) Turn the TILT control to the full tilt UP position.

(d) Whilst taxying, and when clear of personnel, buildings, other aircraft and fuelinstallations turn the RANGE switch to the 20 nm, 50 nm and 150 nm rangepositions, checking for timebase sweep, cloud returns (if appropriate) and rangemarker illumination. Return the RANGE switch to the STANDBY position.

(e) When carrying out the above check the absence of any ground or cloud returns on acloudy day, will obviously be indicative of a system failure. Similarly any spoking ofthe radar (radial lines on the AWR screen eminating from the radar origin like thespokes of a wheel) will almost certainly be caused by a fault within the radar system.

(f) Before take-off, turn the POWER switch to the ON - STAB ON position, ensure thatthe function switch is in the WEA mode and set the tilt control as required.

(g) When airborne and clear of the ground with a positive rate of climb select theappropriate range.

(h) It should be noted that when carrying out a functional check on a coloured screenAWR there is a specific test facility the function of which is described towards the endthis chapter.





Determining the Height of Cloud Tops20. The approximate height of the tops of active clouds may be determined by using the weatherradar in the WEA mode with the stabilisation circuit ON. The angle of tilt of the scanner isgradually increased until the cloud in question just ceases to paint on the screen. The tilt elevationand the range of the cloud are then noted. Figure 10-8 illustrates the situation when the pencil beamis just clearing the top of the cloud. The tilt angle shown on the tilt control is appropriate to thecentre of the pencil beam, and so to determine the elevation angle of the lower edge of the beam it isnecessary to subtract half the beam width. The height of the cloud above the aircraft is nowcalculated using trigonometry, and converting range in nautical miles to feet. Evolving the formulagives:

Height of cloud top above the aircraft (feet)

= Tan (tilt angle - half beam width) x range in feet

= Tan (tilt angle - half beam width) x 6080 x range in nm





FIGURE 10-8Calculation of Cloud Height

21. Taking the example given in Figure 10-8, where the cloud is at 45 nm range, the aircraft is at6000 feet amsl, the tilt angle at which the cloud just disappears is 5 degrees, and the beam width is 4degrees.

Height of cloud top above the aircraft (feet) = Tan (5° - 2°) x 6080 x 45 nm

= Tan 3° x 6080 x 45

= 14,230 feet

The top of the cloud is therefore at 14,230 + 6000 = 20,230 ft amsl.





In the event that you are not given/aware of the value of the tangent of the angle (tilt angle - halfbeam width), the 1 in 60 rule gives the following easier formula:

In the example previously considered the cloud tops would appear to be at 19,500 ft using thisformula.

Using a Monochrome AWR for Weather Avoidance22. As discussed, the contour facility on a monochrome weather radar is a very useful tool indistinguishing between moderate and heavy target returns. The problem is that the ‘hole’ whichindicates a heavy return is the same colour as those areas where the level of target return is either nonexistent or so low as to not paint at all on the screen. The problem is illustrated at Figure 10-9 andFigure 10-10. At Figure 10-9 cumiliform cloud is shown on the 50 nm range with the contour off(WEA is selected on the function switch). At Figure 10-10 the same cloud is shown painting with thecontour on (CONT is selected on the function switch).

Height of cloud top above the aircraft (feet) = (tilt angle - half beam width) x 100 x range in nm





FIGURE 10-9Weather Radar Screen Display with ‘WEA’ Selected





FIGURE 10-10The Effect of Selecting ‘Cont’ Function

23. Were the operator to use the radar continuously in the contour mode it would be easy tointerprete the ‘apparently weak’ paint of the weather ahead as cloud of little significance. It istherefore recommended that the operator alternates between the WEA and the CONT functionswhen assessing the severity of weather returns and planning the subsequent path for weatheravoidance. Areas which paint in the WEA mode but not in the CONT mode should most definitelybe avoided. With some monochrome systems this is done automatically when the contour mode isselected. This gives the ‘holes’ a flashing appearance which serves as an attention getter.





24. Another problem which needs to be adressed is that of the signal strength contour gradient.We know that a zero paint area which is in fact a hole represents an area of large water droplets; thatan area which is painting (in the contour mode) represents an area of smaller but still significantwater droplets; and that the no paint area outside this represents an area of little or no signal return.Whilst it is prudent to associate the ‘hole’ with moderate or severe turbulence, the worst turbulencemay in fact be encountered where the signal gradient is steepest, in other words where the size of thedroplets is changing very rapidly. Such areas are indicated on the screen as a narrow paint betweenthe ‘hole’ and the free air outside of the cumiliform cloud, as indicated at Figure 10-11.

25. Finally, it will be apparent that the setting on the tilt control will have a major effect on theuse of an AWR for weather avoidance. With the tilt set too far downwards the pilot could spend aconsiderable amount of time ‘avoiding’ clouds that are below the aircraft flight path. Similarly, withthe tilt set at too high an angle the presence of a thunderstorm may not be detected until it is too lateto take avoiding action.





FIGURE 10-11Identification of Turbulent Area using an AWR Display

I

Use of AWR for Navigation Position Fixing26. Study Figure 10-12 which shows an AWR in ground mapping mode with the tip of apeninsular of land showing on the screen. At the time of this observation the heading of the aircraft is338°C, the compass deviation on this heading is 2°E, and the aircraft is at a position where thevariation is 11°E. The pilot now wishes to use the AWR information to plot his position on a chart.





FIGURE 10-12Use of AWR for Navigation Position Fixing

27. An examination of Figure 10-12 reveals that this particular radar scans from 90° left to 90°right of the nose of the aircraft therefore the four etched bearing lines represent 30° and 60° right andleft of the nose. The tip of land displayed on the screen is therefore presently 30° right of the nose, oron a bearing of 030° relative. In the same way the range markers indicate that this must be a 50 nmdisplay and therefore the tip of land is at a range of 45 nm.

28. In order to plot the aircraft position on a chart the pilot would have to make the followingcalculations:

aircraft heading 338°C

compass deviation 2°E





29. The bearing of the tip of land is 030° relative and therefore the bearing to plot is calculated inthe same manner as a bearing from an RBI :

30. The pilot would therefore plot the aircraft position as 201° T/45 nm from the fixing point.

Coloured Screen Weather Radars31. As already mentioned, coloured screens use green, yellow and red as the basic colours toshow the ascending strengths of the signal returns.

32. The control and display unit of a colour weather radar is shown at Figure 10-13.

33. The Display Select buttons enable the operator to limit the scan of the dish aerial to one ofthree sectors (left, right or ahead) and also to freeze the screen display.

magnetic heading 340°M

magnetic variation 11°E

true heading 351°T

true heading = 351°

relative bearing = 030°

bearing TO tip of land = 381° - 360° = 021°T

bearing FROM tip of land = 021° + 180° = 201°T





34. Automatic gain control operates in all modes other than the mapping mode where manualgain is available. The intensity control should not be confused with the gain control, it simplybrightens or dims the entire screen display to account for differing light levels on the flight deck.

35. The GCS button is functional only in the weather (WX) mode and suppresses ground clutterin order to give a cleaner paint.

36. The test button causes a coloured test pattern to be painted on the screen to ensure that allcolours are available, as well as running a ‘self test’ programme on the antenna elevation and scancircuitry.

37. The contour mode of a monochrome system is replaced by the weather plus turbulence (WX/T) mode on this coloured radar. With this mode selected, areas of great signal strength and also areaswhere there is a steep signal strength gradient, are painted magenta.





FIGURE 10-13Typical Colour Weather Radar

38. With EFIS equipped aeroplanes the coloured weather returns are normally displayed on theelectronic horizontal situation indicator (EHSI) screen rather than on a dedicated weather radarscreen. This makes it much easier to assess the proximity of any weather to the planned track.





Factors Affecting the Range of an AWRHeavy Rain. In the event that there is an area of heavy rain in front of the aircraft, virtually all of thetransmitted energy will be reflected by the water droplets thereby creating a ‘shadow area’ behind therain into which no radio waves will penetrate. Any active clouds within the shadow area maytherefore not show on the radar screen until the aircraft is at a much closer range than normal.

Water in the Antenna Radome. Any deposits of water present in the radome surrounding the aerialwill prevent energy from being transmitted in that particular direction giving a blank spot in theradar.

Ice Accretion on the Radome. Icing on the radome will cause attenuation of the transmitted andreceived signal, such that targets which would have been displayed will remain undetected until veryclose.







A frequency which is commonly used for AWR is ___________, which has a wavelength of__________ and which lies in the _________ band.

QUESTION 2.

Which types of aerial are generally used for AWR equipment?

QUESTION 3.

In AWR what is the name given to the narrow beam produced by a parabolic dish aerial?

QUESTION 4.

Sidelobe energy is often quite useful in an AWR as it can produce a ______________ on the radardisplay:

QUESTION 5.

What does the "POWER ON STAB OFF" position imply when selected on an AWR control panel?

QUESTION 6.

What are the three range scales commonly used on a typical AWR:





QUESTION 7.

Whilst some AWRs scan 90° either side of the aircraft nose, a more typical figure would be:

QUESTION 8.

What are the four modes of operation of an AWR?

QUESTION 9.

In order to use an AWR for long range mapping, the function switch must be selected to_____________ which means that the ___________ beam will be utilised.

QUESTION 10.

In MAP mode are the AGC circuits in operation, or does the operator have manual control of thegain?

QUESTION 11.

When using a monochrome AWR in CONT mode, where would you expect the area of worstturbulence to be when interpreting the display?

QUESTION 12.

A storm echo just ceases to paint on an AWR screen when the tilt control is 4° up. The radarbeamwidth is also 4° and the range of the storm is 50nm. How far above/below the aircraft is the topof the cloud?





QUESTION 13.

When using an AWR for navigation position fixing, the bearing of a particular landmark is 30° leftof the nose and it is at a range of 35nm. Given that the heading of the aircraft is 337°(M) andmagnetic variation is 9°E, what is the bearing to plot on a chart from the specified landmark?

QUESTION 14.

The colours used in a coloured screen AWR, in descending order of severity, are:

QUESTION 15.

What is the purpose of the GCS button in a typical coloured screen AWR:

QUESTION 16.

What is the main effect of Ice Accretion on the radome covering an AWR:

ANSWERS:ANSWER 1.

9375 MHz 3.2 cm SHF

The most commonly used transmitter frequency for airborne weather radar systems is 9375 MHz inthe SHF band. This frequency gives a wavelength of just over three centimetres.

ANSWER 2.

Commonly used AWR aerials are a Parabolic Dish or a Flat Plate Planar Array





ANSWER 3.

The narrow beam produced by the AWR parabolic dish aerial is called a Conical Beam or a PencilBeam

ANSWER 4.

Height ring

ANSWER 5.

It means that any tilt angle setting on the tilt control will be referenced to the aircraft longitudinalaxis

ANSWER 6.

20, 50 and 150 nm

ANSWER 7.

45° - 60° either side of the nose

ANSWER 8.

WEA (weather), CONT (contour), MAP and MAN (manual)

ANSWER 9.

MAN should be selected on the function switch, therefore the conical (pencil) beam will be used.





ANSWER 10.

In MAP mode manual control of the gain is required

ANSWER 11.

The worst areas of turbulence would be indicated by a narrow paint between the "hole" and the freeair outside of the cloud

ANSWER 12.

Height of cloud (above/below aircraft) = (+ 4.0 - 2.0) x 100 x 50 = +10,000 ft, ie. above aircraft

ANSWER 13.

Answer = 136°(T)Heading 346°(T)

Target rel brg 330° 676° - 360° 316°(T)

Plot bearing of 316° - 180° = 136°(T) from target

ANSWER 14.

MAGENTA, RED, YELLOW and GREEN





ANSWER 15.

The GCS button suppresses ground clutter in order to give a cleaner picture

ANSWER 16.

Ice accretion causes attenuation of the transmitted and received signals thereby reducing detectionrange






SSR Frequencies

Ground Antenna

Aircraft Antenna


Mode A Operation

Mode A Validation

Mode C Operation

Mode C Validation

Mode S





11Secondary Surveillance Radar

1. Secondary Surveillance Radar (SSR) is a development of a military equipment which wasintroduced during the last war to positively distinguish between friendly and enemy aircraft (IFF).

2. One of the functions of SSR is to positively identify an aircraft in flight which is painting as atarget on the radar screen. SSR therefore eliminates the possibility of the mis-identification of anaircraft, which is a distinct possibility if an identification turn, or the pilot's estimation of hisaircraft's position, is used to identify a target on the radar screen in an area of congested traffic.

3. The basic target paint on the screen is termed the primary return, logically since it is theproduct of the primary radar equipment. Assuming that both the ground radar station and theaircraft are SSR equipped a four-figure code will appear on the screen adjacent to the primary paint,and this secondary return will correspond to the code selected by the pilot on the airborne SSRequipment.

4. A second function of SSR is to present the controller with a continuous readout of theaircraft's height, normally presented as pressure altitude. Now the primary target paint isaccompanied by the four-figure code previously discussed, plus another three-figure group denotingthe aircraft's flight level.





5. A PPI radar screen showing both primary and secondary target information is shown atFigure 11-1. The airways shown on the screen are electronically produced. Reporting points, majorairfields, danger areas and Altimeter Setting Region (ASR) boundaries are normally alsosuperimposed, but these have been omitted for clarity. Note the primary paints with their distinctivetails indicating the direction of travel of the aircraft, together with their associated SSR codes andflight level read-outs.

6. A further stage of computerisation at the ground station enables the controller to replace SSRcodes with the aircraft's callsign and supplementary information may follow the height read-out toindicate the aircraft type, destination, intended routing, or ground speed. An example of a typicaltarget display as seen on the air traffic controller’s radar is shown at Figure 11-2.





FIGURE 11-1Typical PPI Radar Screen





FIGURE 11-2Typical Target Display





SSR Frequencies7. The ground-based interrogator transmits on 1030 MHz and receives on 1090 MHz. Theairborne transponder transmits on 1090 MHz and receives on 1030 MHz. Both of these frequenciesare within the UHF band and so the maximum theoretical range of the system is limited to line ofsight.

Ground Antenna8. SSR is a secondary radar system operating, as it were, the other way round to DME. In otherwords with SSR the interrogating radar is ground based and the transponder is airborne. Theinterrogator transmits uni-directionally, the SSR aerial being located on the top of the primary radarhead, or separately as shown at Figure 11-3.





FIGURE 11-3Typical SSR Antenna (dimensions approx 4m x 1.5m)

9. The horizontal polar diagram radiated by an SSR aerial is very narrow, as depicted at Figure11-4, and will invariably have several sidelobes of energy eminating from the aerial but in differentdirections to the main beam.





FIGURE 11-4SSR Antenna - Horizontal Polar Diagram

10. The vertical polar diagram is very broad but radiates minimal output below the horizontal orat very high angles of elevation (i.e. most of the radiation is directed at those angles of elevationnormally used).

Side Lobe Suppression (SLS)11. When a reply is received its angular position on the controller’s PPI is determined by thedirection of the main lobe radiation from the interrogating aerial. If the reply is due to aninterrogation from a side lobe then the indicated bearing will be incorrect. The most commonmethod of suppressing replies to side lobe interrogations is described in the following paragraph.





12. SSR transmits a series of interrogation pulse pairs (see later), the individual pulses of whichare separated by a specific time interval, from the directional antenna previously described. A thirdpulse, called the Side Lobe Suppression Control Pulse (SLS), is radiated from an omnidirectionalantenna (see Figure 11-5) at a time between the original pulses. The amplitude of the SLS controlpulse is such that, if an aircraft were positioned in the main lobe of the SSR, the two interrogationpulses would be received more strongly than the SLS control pulse and an aircraft response wouldtherefore be generated. If the aircraft was positioned outside the main lobe, the SLS control pulsewould be the strongest of the 3 pulses and no response would be generated.

FIGURE 11-5SLS Control Pulse - Polar Diagram





Aircraft Antenna13. It should be noted that the transmitter polar diagram for the aircraft antenna is circular (i.e. ittransmits response pulses omnidirectionally).

Principle of Operation14. Since the SSR transmitter has only to transmit enough energy for a one-way journey, itspower output is much lower than the output of the associated primary radar. The interrogatortransmits a pair of pulses as an interrogation signal, and it is the spacing between the leading edges ofa pair of pulses which determines the Mode of interrogation. There are presently four modes:

Mode A - The interrogator transmits two pulses 8 microseconds apart.

Mode B - The interrogator transmits two pulses 17 microseconds apart.

Mode C - The interrogator transmits two pulses 21 microseconds apart.

Mode D - The interrogator transmits two pulses 25 microseconds apart.

Mode A - Achieves positive identification of primary radar returns, using the four-figure coding already discussed.

Mode B - Serves the same purpose as mode A, but is not presently used in Europe

Mode C - Achieves height (pressure altitude) readout, using the three-figure coding already discussed.

Mode D - Presently used to research future possible applications of SSR.





15. Current airborne SSR transponders normally operate only on Mode A and Mode C, as shownat Figure 11-6, where the ON position on the function switch gives Mode A operation only, and theALT position gives both Mode A and Mode C operation.

16. A further mode, known as Mode 'S' is designed to provide an encoded data link between aground station and an aircraft, or between one aircraft and another aircraft. Mode S is discussed atthe end of the chapter.

FIGURE 11-6Typical Airborne SSR Control Panel





Mode A Operation17. As already discussed, if the airborne transponder is being interrogated for Mode Ainformation the coded interrogation signal will consist of pulse pairs, 8 microseconds apart. In replyto this interrogation, the transponder will transmit omni-directionally a series of up to 14 pulses.The first and last pulses, the frame pulses, are always transmitted. By either including or omittingany or all of the twelve pulses in between the two frame pulses, 4096 possible combinations ofunique response codes are obtained.

18. Any of these codes (response pulse combinations) may be selected by dialling the appropriatenumber between 0000 and 7777 (which will not contain the figures 8 or 9) in the code selectedwindows of the transponder.

19. The coded pulse train transmitted by the transponder in response to a Mode A interrogationwill arrive back at the ground station. Here a computer decodes the signal and paints theappropriate four-figure code (corresponding to the figures selected at the transponder) on the radarscreen adjacent to the primary radar paint.

20. Figure 11-7 shows the reply pulse train with all twelve pulses transmitted. The letters A, B, Cand D refer to the four Mode A code numerals in sequential order. The numerical values assigned toeach pulse are summated by the ground equipment to decode the transmitted pulse train. It will beseen that each numeral has three pulses assigned to it, with values of 1, 2 and 4. If all pulses aretransmitted, as in Figure 11-7, the code selected must be 7777.





FIGURE 11-7Total SSR Pulse Profile

21. Figure 11-8 shows how some pulses would be suppressed in order to transmit the code 5432

FIGURE 11-8SSR Pulse Profile for Code 5432

22. A further pulse (SPI - Special Position Identification) is transmitted 4.35 microseconds afterthe second frame pulse whenever the pilot presses the IDENT button on the transponder controlunit. This ident pulse will be continuously transmitted for 20 seconds once the button is depressed,and the indication to the radar controller when this happens is, typically, that a ring appears on thescreen encircling the primary return.





23. The pilot normally selects a code at the transponder only when requested to do so by airtraffic control. There are however occasions when the pilot should automatically select certaincodes, as follows:

(a) Code A7700 is selected to indicate an emergency situation.

(b) Code A7600 is selected to indicate a radio failure.

(c) Code A7500 is selected to indicate an unlawful interference with the flight (hijack).

(d) Code A2000 is selected when crossing a European boundary inbound and noalternative code has been assigned by air traffic control.

(e) Code A7000 is selected as a conspicuity code, to be used at all times within the UKFIR/UIR by SSR equipped aircraft except when;

(i) discrete Mode A code has been assigned by air traffic control.

or:

(ii) one of the other special purpose codes is being used (2000/7500/7600/7700).

(f) Code A7007 is selected by aircraft operating under Open Skies Treaty arrangements.

24. The indication to the radar controller when code A7700 is selected by the pilot is, typically,that a flashing SOS appears on the screen adjacent to the primary paint. For A7600, a flashing RTFappears, and for A7500 a flashing HIJ.





Mode A Validation25. A controller assigning any Mode A code must validate the code by checking as soon aspossible, either by direct reference to his display or with the assistance of another controlling agency,that the data displayed corresponds to the code which has been assigned.

26. The Mode A code must be checked by one of the following methods:

(a) Instructing the aircraft to squawk the assigned code and observing that the correctnumbers appear on the radar display.

(b) Instructing the aircraft to squawk IDENT and simultaneously checking the screen.

(c) Matching a radar return already identified by primary radar with the assigned codefor the flight.

(d) When an aircraft which is squawking a Mode A code which has previously beenvalidated is handed over from one radar unit to another, the unit accepting the aircraftmay also accept that the Mode A readout is validated.

27. Additionally, at units where code to callsign conversion equipment is in use, procedures toensure the correct correlation of the callsign to the assigned code must be applied.

28. If the Mode A code readout does not correspond to that assigned, the pilot is instructed to re-cycle the assigned code. If this fails to achieve display of the assigned code, the pilot is theninstructed to select code A0000. If a corrupt code still exists the pilot will normally be instructed toswitch off the transponder, however the controller may under certain circumstances elect to use thecorrupt code in order to assist identification and tracking.





29. From the pilots point of view, the important thing is that the unserviceable SSR transpondershould be rectified on landing. If this is not possible and the next sector is one on which the carriageof SSR equipment is mandatory, the permission of ATC must be sought to continue with the nextsector, and the flight plan must be annotated accordingly (even if it is a pre-stored flight plan for aregular schedule).

Mode C Operation30. The transponder is interrogated for Mode C information by firing a series of pulse pairs 21microseconds apart from the ground station. The transponder now produces one of the 4096 codesregardless of the code selected in the window. The code produced is determined by the output of theheight encoder of a pressure altimeter. This output will always be based on a 1013.2 mb datum andis quite independent of any sub-scale setting. Automatic altitude telemetering is available up to128,000 feet. Altitude is displayed in 100 ft increments.

31. The Mode C response pulse train is again transmitted omni-directionally from the aircraftand is received at the radar head on the ground. This time the computer will decode the pulse trainand paint the appropriate flight level on the screen (together with the Mode A code) adjacent to theprimary target paint.

32. With many modern ground SSR systems, it is possible to manually input into the radarcomputer the current QNH. For aircraft flying below the transition altitude, the computer will nowconvert the flight level decode into altitude, and the three-figure height readout will then be replacedby a two-figure readout followed by the letter Z or A on the screen. For example 15Z will indicateto the controller that the aircraft is flying at 1500 feet relative to the QNH.





Mode C Validation33. Before the radar controller uses the Mode C readout to effect vertical separation of aircraft,the Mode C readout must also be validated. This is achieved by asking the pilot to state his flightlevel. Providing that the read-out on the screen is within ± 200 feet of the stated flight level, the

Mode C read-out is thereafter considered to be within acceptable limits. Whilst a tolerance of ± 200

feet is used within the UK FIR/UIRs, in other parts of the world the ICAO tolerance of ± 300 feetmay be accepted.

34. If the Mode C output is found to be in error, the pilot will normally be asked to switch off theMode C, and to continue with Mode A only. If independent switching of Mode C is not possible, thepilot may be asked to squawk code A0000, and to continue with this code in order to indicate thatthe associated Mode C read-out is corrupt.

35. Problems associated with the Mode A/C system are:

• Garbling, ie; interference due to overlapping replies from two or more aircraft in close prox-imity in azimuth and distance. The target aircraft have to be less than 1.7 nm apart (10,000ft) measured in the vertical plane perpendicular to, and from, the ground antenna.

• Fruiting, ie: interference at one interrogator caused by the replies from a transponder inresponse to interrogations from another interrogator.

• Availability of only 4096 codes in Mode A for the identification of aircraft.

• Shielding of the antenna caused by the attitude of the aircraft.

• Additionally the Mode C capability is limited to identifying altitudes by 100 ft increments.





Mode S36. The allocation of 24 bits for aircraft address in Mode S gives it a considerably greatercapacity than Mode C; specifically, it permits individual communication with over 16 million aircraft(sufficient for all aircraft currently flying throughout the world). Mode S differs from Modes A, Band C in that it can function as a communcation system as well as a secondary radar surveillancesystem. This is known as data linking and has the potential to drastically reduce the volume ofverbal information which presently passes between the aircraft and the ground by means ofcongested R/T channels. Finally, Mode S offers the facility to identify altitudes in increments of 25 ft.

37. Mode S can work in any of the following ways:

Selective Addressing. The Mode S equipment addresses a particular message to a specific ModeS address, thereby permitting direct communication between two users.

Mode ‘All Call’. In order to aquire further Mode S equipped aircraft a special roll callinterogation is broadcast at intervals. Any Mode S transponders within range will recognise the rollcall request and will reply with an ‘all call’ reponse consisting of the aircraft identity plus thecapability of the onboard equipment.

Selective Calling. Similar to the above, however only specific Mode S addresses are asked torespond.

Levels of Mode S Transponders38. ICAO Annex 10 stipulates that Mode S transponders shall conform to one of four Levels ofcapability as described below.





European Regulations for the Carriage and Operation of Mode S Airborne Equipment39. Eurocontrol have submitted the following requirements to ICAO as the basis for anamendment to ICAO Doc 7030 to mandate the carriage and operation of Mode S airborneequipment.

(a) For IFR/GAT flights, in airspace designated by the appropriate ATS Authority, a Level2 Mode S transponder as a minimum, with Downlink Aircraft Parameters (DAP)capability (Basic and Enhanced Surveillance Functionality), required by new aircraftwith effect from 1 January 2001 and by all aircraft with effect from 1 January 2003.

Level 1: This is the basic level which permits surveillance based on Mode A/C as well as Mode S. It incorporates a uniquely assigned 24 bit Mode S aircraft address which enables minimum capability for operation with Mode S interrogators. It has no additional data exchange capability and is not prescribed for use on international flights within the European region.

Level 2: Incorporates automatic aircraft identification reporting and standard length air/ground and ground/air data exchange in addition to the Level 1 capability. It is the minimum level permitted for international flight.

Level 3: Incorporates the Level 2 capability with the addition of uplink (ground/air) extended datalink communications.

Level 4: Incorporates the Level 3 capability but allows extended downlink (air/ground) datalink communications.





(b) For VFR flights, conducted in Class B and C Airspace as designated by the appropriateATS Authority and in defined portions of Class D, E, F and G Airspace where thecarriage and operation of SSR transponders has already been prescribed, a Level 2transponder as a minimum, with DAP capability (Basic Functionality), required bynew aircraft with effect from 1 January 2003 and by all aircraft with effect from 1January 2005.

(c) Mode S equipped aircraft shall report automatically Basic Functionality DAPs whichincludes aircraft identification (callsign used in flight).

(d) Mode S equipped aircraft with a maximum mass in excess of 5700kg or a maximumcruising true airspeed in excess of 176 kt (324 Km/h) shall operate with antennadiversity (subject to airframe practicability).

(e) Specific provisions relating to State aircraft shall be subject to regulations issued bythe States concerned.

Downlink Aircraft Parameters40. The specific requirements for DAPs are classed separately as follows :

(a) Basic Functionality:

Automatic Reporting of Flight Identity (callsign used in flight);Transponder Capability Reporting;Altitude Reporting in 25ft increments (subject to aircraft capability).Flight Status (airborne/on the ground).





(b) Enhanced Surveillance Functionality:Basic functionality with the addition of:

Magnetic heading;Speed (IAS/TAS/Mach No);Roll Angle;Track Angle Rate;Vertical Rate (barometric rate of climb/descent, or, preferably baro-inertial);True Track Angle/Ground Speed.

(c) Intended Future Use Functionality:

Additional DAPs which include those relating to aircraft intention are currently underevaluation. Their employment in Mode S Enhanced Surveillance is subject to theresolution of certain technical and institutional issues. In addition, it is anticipatedthat an extended squitter capability and Surveillance Identifier (SI) functionality willbe required following formal adoption of ICAO SARPS.

NOTE:

Basic Functionality DAPs are defined in ICAO Annex 10/Manual on Mode SSpecific Services (Doc 9688).





NOTE:

Enhanced Surveillance DAP parameters are defined in the ICAO Manual onMode S Specific Services, BDS Registers 5,0 and 6,0.

NOTE:

Squitter is the ability of a transponder to automatically transmit pre-formattedinformation which is not in response to an interrogation request.

41. Despite its sophistication, Mode S operates on the same basic principle as conventional SSRand employs the same interrogation and response frequencies.

Future Expansion of Mode S Surveillance Services42. In anticipation of further expansion of Mode S Surveillance Services, consideration has beengiven to the downlinking of additional aircraft parameters. Those which indicate aircraft intention(selected parameters) offer the greatest potential benefit to the ATM system and in particular to thesafety nets in terms of enhanced tracking and anticipated knowledge of aircraft manoeuvres.However, the resolution of certain technical and institutional issues associated with the downlinkingof these parameters is essential before they can be introduced for operational use. Therefore, oncethese issues have been resolved, the following parameters as defined in the ICAO Manual on Mode SSpecific Services, BDS Register 4,0, are likely to be recommended for inclusion in regulations:

• Selected Flight Level/Altitude





• Selected Magnetic Heading

• Selected Course

• Selected IAS/Mach No

Future Airborne Datalink Equipment Requirements43. It is important that aircraft operators are made aware as early as possible of planned groundinfrastructure developments which could affect future airborne equipment carriage requirements.Mode S Level 2 transponders, as a minimum, have been prescribed because the employment of thefull Mode S datalink has not yet been endorsed in the context of an overall European DatalinkStrategy. However, work is ongoing to evaluate its suitability for this purpose. Endorsement of theextended use of Mode S datalink would lead to a requirement for Mode S Level 4 transponders as aminimum.

Airborne Collision Avoidance Systems (ACAS)44. ACAS SARPS were adopted by ICAO in 1995. However, since December 1993, the TrafficAlert and Collision Avoidance System (TCAS II) has been mandated for use in US Airspace byaircraft of more than 30 passenger seats. Additionally, evaluation of TCAS II operations has beenongoing in European airspace for a number of years. These systems, which incorporate the use of aMode S transponder integral to the TCAS system, interrogate both Mode S and Mode A/Ctransponders of other aircraft. The received responses are processed to provide, where appropriate,collision avoidance in the vertical plane of traffic in the vicinity of ACAS equipped aircraft.





45. European ACAS policy is to require the mandatory carriage and operation of an ACAS IIconforming to ICAO SARPS in the airspace of ECAC Member States. An implementation schedulehas been adopted, in principle, as follows:

(a) With effect from 1 January 2000, all civil fixed wing turbine engine aircraft having amaximum take-off mass exceeding 15000 kg or maximum approved passenger seatingconfiguration of more than 30 will be required to be equipped with ACAS II.

(b) With effect from 1 January 2005, all civil fixed wing turbine engine aircraft having amaximum take off mass exceeding 5700 kg or maximum approved seatingconfiguration of more than 19 will be required to be equiped with ACAS II. It shouldbe noted that the weight and seat parameters are subject to confirmation.






Flight Management Systems

Boeing 737-400 FMS Operation

Flight Director Systems

Electronic Display Systems

Electronic Flight Instrument Systems (EFIS)

VOR/DME Area Navigation (RNAV)





12Area Navigation Systems

1. Area navigation (RNAV) is the process of calculating the range and bearing from the aircraftof any given location within the area, in order to fly a straight-line track to that point. It isparticularly useful when the desired geographical location is not marked by any radio navaid, as isoften the case when operating off airways. In many parts of the world RNAV routes have beendeveloped to allow navigation outside standard routes thereby decreasing traffic congestion andmaking optimum use of the available airspace.

2. All RNAV systems are computer based. Complex RNAV systems may use inputs from allavailable navaids (INS, Loran, VOR, DME) plus TAS, altitude and heading inputs from the CentralAir Data Computer (CADC). More basic systems give range and bearing to a selected location usingexternal inputs from VOR/DME only, with internal inputs of latitude/longitude of the selectedlocation and of the VOR/DME stations used to achieve the solution. The latter would normally bepre-programmed in the computer memory against the VOR frequency.

3. There are two levels of accuracy of operation for RNAV equipment: B-RNAV and P-RNAV.To be eligible for B-RNAV operations , on board navigation equipment will be required to provide

en-route track keeping accuracy of ± 5 nm or better for 95% of the flight time. B-RNAV has anaccuracy comparable with that of aircraft currently operating the present system on routes definedby VOR/DME. Precision RNAV (P-RNAV) requires a track-keeping accuracy of 0.5 nm standarddeviation or better.





4. The student will by now appreciate that, as a general statement, fixing of an aircraft positionusing radio navigation systems (e.g. VOR, DME, Loran, and GPS (see Chapter 17)), will result in abounded error, whilst fixing of an aircraft position using an INS or IRS will result in an unboundederror (i.e. the error gets worse with time). In sophisticated RNAV systems the radio navigationsystem inputs can be used to ‘tie down’ the INS/IRS position information in those areas whereground/space-based fixing cover is good.

5. An RNAV system is programmed to calculate the most accurate continuously updatedposition possible by using the various radio navigation inputs however, should these not be available,it will continue in dead reckoning mode until such information is restored.

6. A general block schematic diagram of an area navigation system is shown in Figure 12-1.





FIGURE 12-1Simple Area Navigation System





7. The purpose of the components of this simple system are briefly described below.

Flight Data Storage Unit This part of the system is where ‘permanent’ information is stored; forexample location, elevation and frequency of beacons and airports, standard departure and arrivalroutes etc.

Automatic Data Entry Unit. This unit allows an operator to automatically feed into the systemthe company’s standard routes together with the relevant waypoint data.

Navigation Computer Unit. The NCU processes the information from the various sensors and,by comparing the result with the selected flight profile, it generates various output commands.

Control Display Unit. The CDU is the interface between the pilot and the NCU and as such itallows the pilot to modify the flight profile, as necessary, and to display selected information in thecockpit.

Compass System. Magnetic heading is fed from the main compass and is fed as an input into theNCU.

Air Data Computer An Air Data Computer provides TAS and altitude inputs to the NCU.

Information Displays Navigational outputs from the NCU are fed to various cockpitinstruments such as the RMI (discussed in Chapter 4) or an EHSI (see later).

Sensor Inputs. Figure 12-1 shows some of the possible sensor inputs into the NCU (preciseavailability will depend on the individual aircraft fit). It should be noted that there are different typesof sensor input :





In large aircraft like the Boeing 737, 747, 757 etc, the tasks associated with Area Navigation will beexecuted by the Flight Management and Guidance System (FMS).

(a) Self Contained on-board systems

INS/IRS/Doppler Present position, as displayed on the CDU of self-contained navigation systems, is fed into the area navigation system as an input (whether it be in geographic co-ordinates or graphical form).

(b) External Sensor Systems Position fixing obtained by using radio navigation aids.

GPS-Position information in latitude and longitude / velocity

DME/DME-Range/Range (RHO/RHO)

DME/VOR-Range/Bearing (RHO/THETA)

(c) Air Data Inputs

TAS typically from a CADC

Altitude

(d) Compass Input

Magnetic Heading





Flight Management Systems8. The Flight Management System (FMS) is an integration of the aircraft subsystems, thepurpose of which is to assist the flight crew in controlling and managing the flight path of theaircraft. The flight path is divided into lateral and vertical profiles, commonly known as LNAV andVNAV. The system allows the pilots to select the degree of automation required at all stages of flightand consequently the need for many routine tasks and computations is eliminated.

9. Primarily the FMS provides automatic three-dimensional navigation, fuel management andfuel monitoring together with the optimising of aircraft performance. It also provides information tothe appropriate displays, including the electronic map, which is fully described in the section dealingwith Flight Directors and Electronic Flight Information Systems (EFIS). FMS also provides airspeedand engine thrust cues.

10. The main components of an FMS are :

(a) Flight Management and Guidance Computer (FMC)

- uses both manual and automatic inputs of data to compute 3 dimensional position, performance data etc in order to fly the aircraft accurately and efficiently along a pre-defined route.

(b) Multipurpose Control and Display Unit (MCDU)

- the interface between the pilots and FMC.

(c) Flight Control Unit





The Flight Management and Guidance Computer11. A schematic diagram of the component parts of a typical flight management system is shownat Figure 12-2. The heart of the system is the Flight Management Computer (FMC) and its associatedMultipurpose Control and Display Unit (MCDU). A CDU of the type found in the Boeing 737 isillustrated at Figure 12-3.

- supplies the commands to control the lateral and vertical flight path of the aircraft.

(d) Flight Management Source Selector

- selects the sources of input to be used by the FMC.

(e) Display System

- any means of displaying the required data/ information to the pilots.





FIGURE 12-2A Typical FMS





FIGURE 12-3A Typical FMS Multipurpose Control and Display Unit (MCDU)





12. The MCDU combines flight plan information entered by the pilots with information suppliedfrom supporting systems and information contained in memory. This enables the FMC to determinethe aircraft position and to provide pitch, roll and thrust information in order to fly the profilerequired. Commands are sent by the FMC to the autopilot, the flight director and the autothrottle(autothrust) system. FMC navigational and performance computations are displayed on the MCDUsfor reference or monitoring. Related FMC commands for lateral and vertical navigation are coupledto the AFDS and Autothrottle through the Mode Control Panel (L NAV and VNAV). The IRSs andother aeroplane sensors provide additional required data. MCDUs also permit interface with theAircraft Communications Addressing and Reporting System (ACARS). Additionally, mapinformation is sent to the Electronic Horizontal Situation Indicator (EHSI) and displayed in themanner described in the section dealing with EFIS.

The FMC Data Base13. The information which is stored in the FMC data base is divided into two main sections,namely navigation information and aircraft performance information.

14. The navigation data includes the location of radio navigation aids, SIDs, STARs, companyroutes, airports, runways, approach aids and airways structures. The data base is tailored to theneeds of the individual carrier. This navigation data base is produced by a specialist agency (such asJeppesen) and is normally updated on a 28 day cycle. Data transfer hardware (using a magnetic tapecassette) is provided to enable the operator to load a new data base into the aircraft FMCs. In orderthat flight operations do not come to a grinding halt at midnight on the last day of validity of theexpiring data base, the current data base together with the next effective data base are both stored inthe FMCs. For the pilot then, step one when setting up the FMCs is to ensure that the correct database for the date of the flight is the operational one.





15. Within a given 28 day period it is likely that certain information contained in the navigationdata base will become invalid, for example NOTAMs may inform us that a given VOR is out ofservice for a period of time. The pilot can access the data base and delete that VOR, but only for theduration of the flight. It is therefore impossible for the pilot to corrupt the data base itself. It isimportant to remember that the data base is produced by another human being and may thereforecontain errors. Because of the high degree of automation involved when, basically, the FMC isdriving the aeroplane, it is essential that the pilots monitor the aircraft’s progress using conventionalnavigation techniques (raw data), and also that any errors in the data base are fed back throughreporting channels so that they can be remedied.

16. During flight the FMC will search the navigation data base and automatically select the besttwo DME stations with which to determine the aircraft’s present position. In the absence of suitableDME/DME crosscuts the system will use co-located VORs and DMEs. When DME/DME or DME/VOR fixing is not possible, for example on an oceanic leg, the aircraft’s position is determined by theinertial reference systems plus a correction vector that has been developed by a Kalman filter over aperiod of time. In those systems that use GPS position as an input into the FMC, it is usually possiblefor the pilot to delete any satellite that has automatically been selected by the GPS receiver, in orderto obtain the best fix geometry.

17. The Kalman filter uses hybrid navigation techniques. It takes, for example, positioninformation from a number of sources and then statistically analyses that data (taking into accountthe possible errors) to produce a final solution which, in the case of position, would be the FMCposition. The filter also produces the correction vector discussed in paragraph 16.

18. Take the situation where an aircraft, equipped with say 3 inertial systems, is flying fromEurope to the USA. As the aircraft crosses the UK, on its way to join the NAT track system, the FMSwill be using DME/DME radio ranges to assist in determining position. Figure 12-4 gives a pictorialpresentation of the computations involved.





FIGURE 12-4Position Determination by an FMC

19. In simple terms the FMC first averages out the 3 IRS positions to determine a ‘mean’ inertialposition. Secondly, it compares the mean inertial position with the radio aid position (in this caseDME/DME ranging is used) and, taking account of the likely error in each position, it computes afinal FMC position which is used to steer the aircraft along the planned track.

20. The position correction vector in the above example stretches between the mean inertialposition and the final FMC computed position. (In an aircraft equipped with a single inertial systemthe vector would obviously start from that single position).





21. It will be obvious from the above explanation that, in order to develop the position correctionvector over a given time, there must be a continuous supply of radio information. However, once theaircraft leaves the area of ground based radio aids the FMC can still use the ‘history’ of the vector todevelop it further, and hence continues to provide the best possible estimate of position. As theaircraft coasts in again over the USA radio aid fixing will once again be used to ‘tie down’ the FMCposition.

22. The accuracy of a Kalman filtering system such as the one described is dependant upon twomain factors :

(a) The quality and complexity of the Kalman filter design.

(b) The error characteristics of the various ‘navigation’ sensors used by the system mustbe complementary. (i.e. any single system input which is subject to alot of ‘noise/variation’, or ‘drifts’ in value, may cause a significant error in FMC computedposition).

23. The FMCs will automatically select the VOR/DME stations which are displayed on the EHSIneedles, the standby RMI needles and the DME range readouts. The system will decode the morseidentifier and display letters on the screen. If a satisfactory identifier decode is not achieved, thefrequency will be displayed rather than the identifier. In this event it is up to the pilot to identifyground station in the conventional manner. Similarly, providing that the FMC has been informed thatthe intention is to fly an ILS approach to a given runway at the destination/alternate aerodrome, the





relevant ILS will be autotuned and identified, again with the morse identifier displayed to the pilot,but this time on the Electronic Attitude Direction Indicator (EADI). Where the departure is from anILS runway, the FMC will again autotune the ILS in order to provide centre line guidanceimmediately after take-off. When NDBs form part of a SID, STAR approach procedure or (unusuallythese days) an airways structure, these are also autotuned and identified by the FMC. The optionalways exists for the pilot to override the automatics by ‘hard tuning’ stations of his or her choice.

24. The performance data base contains all of the information normally contained within theperformance manual, such as engine characteristics, the aircraft limiting speeds for the variousconfigurations, optimum/maximum cruise altitudes and an aerodynamic model of the aeroplane. Thedata base may be individually tailored for an individual aeroplane within a fleet. Variables such asfuel quantity, zero fuel weight and a company cost index are entered by the flight crew. This data ispeculiar to the next sector only and is automatically dumped by the FMC following the next landingand engine shutdown. The simplest explanation of the cost index is that it is a numerical value whichtells the FMC whether the operator considers that fuel economy (with larger sector times) orminimum sector times (with a resultant higher fuel burn) is the preferred option. The cost index cantherefore be altered on a sector by sector basis to account for the circumstances of that flight.

Modes of Operation for Dual FMC Installations25. FMC systems are normally duplicated and each FMC has its own CDU. There are 4 modesof operation :





Dual Mode. With the system operating normally the two CDU/FMCs are interconnected andpilot entered data which is entered at one CDU is automatically transferred to the other one. In otherwords, one FMC provides the master function and the other the slave function. The pilots may selecttheir own EHSI display (full or expanded VOR, full or expanded ILS, map or plan) regardless ofwhat is displayed on the other EHSI.

Independent. The first stage of degradation of the system occurs when a disparity is sensedbetween the outputs of the two FMCs. Now each CDU/FMC works Independently of the other andthe pilots are left to identify the serviceable system. Each CDU will supply its own EHSI, howevernow the pictures on each of the EHSIs (assuming that they are in the same mode with the same rangeoption selected) will differ.

Single. The next stage of degradation of the system when one FMC or CDU fails altogether. Youare now down to a single system operation, however both EHSIs can be driven from the same FMC/CDU providing only that both pilots select the same mode and range setting.

Back-Up Navigation. Finally, should both FMC/CDUs fail, the pilots are left with blank EHSIsand the prospect of limited use of the FMS. Navigation is achieved by manually tuning en route andapproach aids which are subsequently displayed on a conventional RMI and analogue DME readout.





Lateral Navigation Guidance26. The FMC calculates the great circle tracks and distances between successive waypoints in theactive flight plan. These are the track lines which are shown on the EHSI map display. The activeflight plan includes the SID, the STAR and any relevant holding patterns. Under normalcircumstances (managed guidance) the FMC will command the autopilot to maintain the definedtrack (at a particular altitude and speed). With the aircraft flown manually (selected guidance) theFMC commands the human pilot to maintain a particular value of a parameter (heading, speed etc)by making selections on the Flight Control Panel (FCP). At any time the pilot can take control of thelateral navigation of the aircraft by going into heading mode. The FMC will automatically revert toheading mode whenever LNAV capture parameters are out of limits or when, for example, awaypoint is reached and no route is defined beyond that point.

Vertical Navigation Guidance27. Providing that the pilot does not modify the climb profile, the FMC will command a climbwith thrust at the airspeed limit associated with the departure airfield until above the speed limitaltitude or flight level. Thereafter the climb will continue at climb thrust and economy speed to thedemanded cruise level. Where altitude/level constraints are imposed by the SID (cross point X at/at orbelow/at or above a given altitude or flight level), these constraints will be shown on the EHSI mapand plan displays. The aircraft will comply with these constraints providing that the FMC remains inthe fully managed mode. In the event that ATC impose an altitude constraint, this can be entered bythe pilot as a vertical revision to the waypoint to which the constraint applies. If, during the climb,the FMC senses that the aircraft will not be able to comply with the constraint due to an insufficientrate of climb, the pilot will be warned. The FMC will capture any altitude which is selected andarmed on the Mode Control Panel (MCP)/Flight Control Unit (FCU).





28. During the cruise, economy speed will be used until the top of descent point.

29. The top of descent point is computed by the FMC, as it were, from touchdown backwards.The FMC has knowledge of the aerodrome elevation, and the QNH is manually entered by thepilots. The exact vertical distance from the cruise level to touch down is therefore known. Flight levelor altitude constraints, as defined by the STAR and the approach procedure are stored in thenavigational data base, and the descent profile is computed to account for these constraints. Thedescent will normally be computed such that, wherever possible, the engines will be at idle power(which is fuel efficient). The descent will be computed at economy speed down to the point wherethe STAR imposes a maximum speed constraint, and thereafter at speeds which will enable the slats/flaps/landing gear to be extended at the appropriate points. Wind velocities for the descent can bemanually entered by the pilots in order to refine the computation.

30. Typical VNAV climb, cruise and descent profiles for a B757 are illustrated at Figure 12-5 andFigure 12-6.





FIGURE 12-5Typical VNAV Climb / Cruise Profile





FIGURE 12-6Typical VNAV Descent Profile





Boeing 737-400 FMS Operation31. The foregoing paragraphs gave a brief overview of an FMS, in general terms. The followingparagraphs describe, in some degree of detail, the operation of the Boeing 737-400 FMS plusinformation regarding the processes involved in providing the system’s navigation solution.

Pre-Flight32. The CDUs are used during preflight to manually initialise the IRSs and FMC with departureinformation such as present position, flight plan routing, zero fuel weight, and planned cruisealtitude. These CDU entries and the data bases then form the starting point for FMC computations.

33. If the permanent data base does not contain all of the required flight plan data, additionalairports, navaids, and waypoints can be defined by the crew and stored in either a supplemental or atemporary navigation data base. Use of these additional data bases provides world-wide navigationalcapability, with the crew manually entering desired data into the FMC via various MCDU pages.Information in the supplemental nav data base is stored indefinitely, requiring specific crew actionfor erasure ; the temporary nav data base is automatically erased at flight completion.

34. Stored waypoint identifiers may be entered manually on either the RTE or RTE LEGS pages,or they may be entered automatically as part of a company route designation. The following arevalid CDU entries for published waypoint indentifiers stored in the permanent navigation data base(five characters maximum) :

• waypoint indentifier (waypoint name)

• navaid indentifier

• runway number





• airport ICAO indentifier

Created Waypoints35. If the permanent nav data base does not contain the desired stored waypoint(s), then new(previously unstored) ‘Created Waypoints’ can be defined by the crew. On the RTE or RTE LEGSpages, Created Waypoints are keyed into the Scratch Pad as any of the following :

• Place Bearing/Distance (for example, SEA250/40), where ‘Place’ is any identi-fier already stored in either the permanent, supplemental, or temporary navdata base.

• Place Bearing/Place Bearing (for example, SEA180/ELN270), the intersectionof bearings from two different ‘Places’.

• Along-Track Displacement (for example, SEA/-10), the distance either side ofan existing flight-plan waypoint.

• Latitude and longitude (for example, N4731.8W12218.3).

36. The waypoints are automatically stored in the temporary nav data base for one flight only.On the NAV DATA pages, entry of the FMC - assigned identifier on the WPT IDENT line provides adisplay of the parameters originally keyed -in to define that waypoint.

37. Alternatively, Created Waypoints can also be initially defined using crew-assigned identifierson either the SUPP NAV DATA or REF NAV DATA pages. This method allows ‘waypoints’ to bedefined in any of three FMC categories ; Waypoints, Navaids or Airports. Entries defined on theSUPP NAV DATA pages (accessible on the ground only) are automatically stored in the supplementalnav data base until deleted by the crew. Entries defined on the REF NAV DATA pages areautomatically stored in the temporary nav data base for one flight only.





38. The supplemental and temporary data bases share storage capacity for forty Navaids and sixAirports, the entries being stored in either data base on a ‘first come, first served’ basis. For theWaypoint category, exclusive storage is reserved in the temporary data base for twenty entries(including those created on the RTE or RTE LEGS pages). An additional twenty Waypoints (up to amaximum of forty) can be stored in either the temporary or supplemental data base on a ‘first come,first served’ basis.

39. When any storage category is full, entries which are no longer required should be deleted bythe crew to make space for additional new entries. Created Waypoints cannot be stored in the database Runway category.

Conditional Waypoints40. The preceding waypoints all refer to geographically-fixed positions. Waypoints which are notgeographically fixed are called Conditional Waypoints, and are embedded within stored proceduresand displayed on the CDU in parenthesis. They cannot be entered manually. Conditional Waypointsare displayed as any of the following:

• (1500) altitude condition

• (SEA330) VOR radial crossing condition

• (SEA-10) DME crossing condition

• (INTC) intercept course to next waypoint

• (VECTOR) maintain heading indefinitely





NOTE:

When (VECTOR) is the active leg, the FMC does not automatically sequenceto the next waypoint. The next waypoint becomes active only uponEXECution of the procedures for Proceeding Direct To a Waypoint orIntercepting a Leg to a Waypoint.

Inertial Reference Systems41. Two independant Inertial Reference Systems (IRSs) are installed, plus Mode Selectors and oneIRS Display Unit (ISDU) located in the cockpit. The IRSs are the aeroplane’s sole source of attitudeand heading information, except for the standby attitude indicator and standby magnetic compass.

42. In their normal navigation mode, the IRSs provide attitude, true and magnetic heading,acceleration, vertical speed, ground speed, track, present position, and wind data to appropriateaeroplane systems. IRS outputs are independant of external navigation aids.

IRS Alignment43. The IRS must be aligned and initialised with the aeroplane position before it can enter theNAV mode. The position is normally entered through the FMC CDU during alignment. If theposition cannot be entered through the FMC CDU, it may be entered through the ISDU keyboard. Atmajor airports present position may be inserted into the CDU by inserting the appropriate Gatenumber (providing the appropriate latitude and longitudes are stored in the database). The aeroplanemust remain stationary during alignment.





44. Normal alignment, between 70°12’ North and 70°12’ South Latitudes, is initiated by rotatingthe IRS Mode Selector from OFF directly to the NAV position. The IRS performs a short DC powertest, during which the ON DC Light illuminates. When the ON DC Light extinguishes and theALIGN Light illuminates, the IRS has begun the alignment process. Aeroplane present positionshould be entered at this time. The IRS will automatically enter the NAV mode after approximately10 minutes, and the ALIGN Light will extinguish.

45. High latitude alignment, at latitudes between 70°12’ and 78°15’, requires an extendedalignment time. The Mode Selector must be left in the ALIGN position for 17 minutes, then rotatedto the NAV position. The IRS will then immediately enter the NAV mode.

46. Magnetic variation between 73° North and 60° South latitudes is stored in each IRS memory.The data corresponding to the present position are combined with true heading to determinemagnetic heading.

Fast Realignment47. During transit stops with brief ground times, a thirty-second realignment and zeroing ofground speed error may be performed by selecting ALIGN from NAV while the aeroplane is parked.Present position should be simultaneously updated by manually entering latitude and longitude priorto reselecting NAV.

NOTE:

If the aeroplane is moved during alignment or fast realignment, (ALIGN lightilluminated), the IRSs automatically begin the full 10-minute alignmentprocess over again.





In-Flight48. The FMC compares planning information with actual data from a number of other sources(including required IRS inputs). With LNAV and VNAV engaged, the CDU displays allow the crew tomonitor proper FMS operation and flight progress. With LNAV and VNAV disengaged, the displaysare used for reference, allowing the crew to fly the selected route/profile either manually or withconventional autoflight modes. The CDUs are also used to : provide ‘what if’ previews of flight planoptions ; make revisions to the flight plan ; and provide reference data.

49. FMC navigational computations are bases upon an ‘FMC position’ which is established usingradio inputs and/or IRS present position. The FMC position may be based upon IRS data only(inertial/dead reckoning mode) ; however, available DME inputs are normally used to refine andupdate the FMC position (radio/inertial mode). Just prior to take-off, the crew may set the FMCposition to a point on the departure runway via the CDU TAKEOFF REF page. Activation of theTO/GA button updates the FMC to this position.

50. It should be noted that radio updating does not occur on the ground. Consequently,navigation position error can accumulate in the FMC during transit. Fast realignment of the IRSswith a new present position removes the errors. The errors will also be removed after take-off whenupdating again becomes available.

51. With normal operation the DMEs are automatically tuned by the FMC. The stations to betuned are selected based upon the best available signals (in terms of geometry and strength) forupdating the FMC position, unless a specific station is required by the flight plan. Radio position isdetermined by the intersection of two DME arcs. Manual selection/deselection of a particular DMEis possible via the MCDU.





52. If the DMEs fail, or if suitable DME stations are not available, FMC navigation is based uponIRS position information only. The two VHF Nav Radios are used by the FMC for LOCALISERupdating during an ILS approach and by the crew for navigation monitoring.

NOTE:

The FMC is designed to automatically reject unreliable navaid data duringFMC position updating. However, in certain conditions, navaids which are inerror may satisfy the ‘reasonableness criteria’ and provide the FMC with aninaccurate radio position. One of the most vulnerable times is when a radioposition update occurs just after take-off. This is usually manifested in anabrupt heading correction after engaging LNAV. The position shift can be seenon the EHSI map which will shift the desired track and runway symbol to aposition significantly different from that displayed during the ground roll.

53. When radio updating is not available, the FMC uses the IRS position as a reference. Thismode of navigation is referred to as IRS NAV ONLY, and a message is displayed to warn the flightcrew that navigation accuracy may be less than required. During IRS NAV ONLY operation, theFMC applies an automatic correction to the IRS position to determine the most probable FMCposition. This correction factor is developed by the FMC by monitoring IRS performance duringperiods of radio updating to determine the IRS error. Flight crews should closely monitor FMCnavigation during periods of IRS NAV ONLY operation especially when approaching thedestination. The accuracy of the FMC navigation should be determined during the descent phase offlight by using radio navaids and radar information if available.





NOTE:

Inaccurate radio updating may cause the FMC to deviate from the desiredtrack.

System Control54. The crew may select any degree of automation desired. This can mean simply using the CDUData Displays for reference during manual flight, or using conventional autopilot functions, orselecting full FMS operation with automatic flight path guidance and performance control.

55. Even with full FMS operation, management and operation of the aeroplane is always underthe total control of the flight crew. The flight crew should monitor FMC navigation throughout theflight to ensure that the desired route is being accurately followed by the automatic systems.





Flight Director Systems56. Flight director systems (FDS) integrate the information presented by traditional flightinstruments (artificial horizon, turn and bank, gyrocompass) with the information received fromexternal sources (VOR, DME, ILS, Radio Altimeter) to produce control commands. This achievesmore accurate flight guidance and control whilst reducing the pilot’s workload in terms ofmonitoring and co-ordinating the many individual sources of information. The FDS presentation isin the form of two displays, an attitude direction indicator (ADI) and a horizontal situation indicator(HSI). The ADI presents flight guidance commands in pitch and roll, whilst the HSI presents thenavigational situation. The two displays are shown at Figure 12-7. Note the command bars on theADI shown at Figure 12-7. An alternative presentation of the command bars are shown on the ADIillustrated at Figure 12-11.

57. The student may have difficulty in visualising the display movement of the ADI inFigure 12-7, hence two diagrams are shown at Figure 12-8 which illustrate the use of the equipment.The first diagram shows commands of 5° pitch up and a right turn; the second diagram shows bothcommands having been satisfied and where 15° of right bank is displayed.

58. Flight director systems have existed for many years as analogue instruments, indeed it is ananalogue system which is shown at Figure 12-7. In this chapter we will subsequently consider asystem which employs modern glass cockpit technology and which therefore presents far moreinformation to the pilots.





FIGURE 12-7‘Conventional’ ADI and HSI Displays





FIGURE 12-8Examples of ADI Indications

Electronic Display Systems59. To display all the necessary information and data concerned with in-flight management of theaircraft systems would demand a vast array of instrumentation, impossible for a typical two or three-person flight-deck crew to comprehensively monitor. Furthermore, much of the data is only relevantat certain flight phases or in particular circumstances and therefore need not be permanentlydisplayed.





60. This has led to the development of electronic display systems (the glass cockpit) in which thedata is processed and stored by large capacity computers and displayed as required on colour CRTscreens in either alphanumeric form or as symbols. The following colours are being recommended inJAR 25 based on current-day common usage. Deviations may be approved with acceptablejustification.

(a) Display features should be colour coded as follows:

(b) Specified display features should be allocated colours from one of the following coloursets:

WarningsFlight envelope and system limitsCautions, abnormal sourcesEarthEngaged modesSkyILS deviation pointerFlight director bar

RedRedAmber/YellowTan/BrownGreenCyan/BlueMagentaMagenta/Green





(c) Precipitation and turbulence areas should be coded as follows:

(d) Background colour: Background colour may be used (Grey or other shade) to enhance display presentation

Fixed reference symbolsCurrent data, valuesArmed modesSelected data, valuesSelected headingActive route/flight plan

Colour Set 1

WhiteWhiteWhiteGreenMagenta**Magenta

Colour Set 2

Yellow*GreenCyanCyanCyanWhite

The extensive use of the colour yellow for other than caution/abnormal information is discouraged. In colour Set 1, magenta is intended to be associated with those analogue parameters that constitute ‘fly to’ or ‘keep centred’ type information.

Precipitation

Turbulence

0 - 1 mm/hr1 - 4 "4 - 12 "12 - 50 "Above 50 "

BlackGreenAmber/YellowRedMagentaWhite or Magenta





61. The screens can be arranged to display primary information on a continuous basis, with faultor emergency information superimposed as necessary. ‘Programmes’ of secondary information can becalled up and displayed when required.

62. The displayed data falls into two broad categories; navigational and aircraft systems.

63. The computer-generated electronic displays which show the navigational data are jointlyknown as the Electronic Flight Instrument System (EFIS). The upper screen shows the ADI whilst thelower screen shows the HSI, either in a format similar to that shown at Figure 12-7 or in one of thepilot selectable formats subsequently discussed.

64. The computer-generated electronic displays which show the aircraft systems are jointlyknown as either the Engine Indicating and Crew Alerting System (EICAS) or as ElectronicCentralised Aircraft Monitoring (ECAM). Basically, EICAS is a Boeing term whereas ECAM is anAirbus term.

65. In the remainder of this section only the EFIS portion of the total electronic display system isconsidered.

Electronic Flight Instrument Systems (EFIS)66. EFIS displays information on two, approximately 5 inch square, screens for each pilot. Onescreen corresponds to the ADI (attitude direction indicator) and the other to the HSI (horizontalsituation indicator), although the computer-generated displays convey far more navigationinformation than is possible with the conventional electro-mechanical flight director system.Figure 12-9 shows a typical interface between EFIS and signal inputs. These displays are capable ofpresenting all of the necessary primary and secondary flight information.





67. Conventional (analogue) back-up pressure and gyro instruments are usually retained forairspeed (ASI), altitude (pressure altimeter), pitch and bank (artificial horizon) and heading (directreading compass).

68. The symbol generators interface between the aircraft systems, the control panels and thedisplay screens. They perform the main control functions of the EFIS, including system monitoringand generation of the digital and analogue displays on the electronic ADI (EADI) and electronic HSI(EHSI) screens.

69. Appreciate that some manufacturers refer to the EADI as the primary display and to the EHSIas the navigation display.





FIGURE 12-9EFIS Interface Diagram





The EFIS Control Panel70. An example of a control panel is at Figure 12-10. Remote light sensors respond to ambientflight-deck lighting levels and adjust the CRT displays accordingly to maintain optimum displayvisibility. Display brightness can also be adjusted manually by brightness controls (BRT) on each halfof the pilots' EFIS control panel (EADI and EHSI). The buttons at the bottom of the panel (the EHSIMap Mode Selector Switches) are illuminated when pressed to select on.





FIGURE 12-10EFIS Control Panel





The EADI Screen71. The upper (EADI) screen conventionally displays aircraft attitude in pitch and roll against ashaded (raster-scanned) background, the upper half of which is coloured blue (cyan) and the lowerhalf yellow (or light brown). The source for the attitude data is the aircraft's inertial referencesystem(s). The EADI also displays flight director command bars for roll and pitch commands, as wellas ILS localiser and glideslope deviation, selected airspeed deviation, ground speed, automatic flightsystem and autothrottle system operating modes, radio altitude and decision height. A typical EADIdisplay is shown at Figure 12-11.

72. Radio altitude is displayed digitally between 2500 feet and 1000 feet agl, as shown in the topright hand corner of the EADI at Figure 12-11. Below 1000 feet agl the display becomes analogue/digital, again as illustrated at Figure 12-11. There is a decision height (DH) setting knob on the EFIScontrol panel (Figure 12-10). At radio altitudes above 1000 feet the selected DH is displayed digitallyon the EADI (Figure 12-11). Below 1000 feet radio altitude the DH is displayed as a magentacoloured marker on the circular analogue radio altimeter scale (Figure 12-11). As the aircraftdescends from 1000 feet radio altitude the white circular scale segments are progressively erased inan anti-clockwise direction, so that the remaining 100 foot segments indicate the height aboveground. At 50 feet above the selected decision height an aural chime alert sounds with increasingfrequency until decision height is reached. The circular scale and marker then both change colour toamber and flash for several seconds. This alert is manually cancelled by pressing a reset button on thecontrol panel (Figure 12-10).





FIGURE 12-11Typical EADI Display





73. Airspeed error above (F) or below (S) a selected airspeed is indicated by a magenta pointerand scale on the left hand side of the EADI. Glideslope deviation is similarly displayed on the righthand side of the screen. Localiser deviation is indicated by a magenta pointer and scale at the bottomof the display. ILS localiser and glideslope deviations are emphasised by the appropriate pointer andscale changing colour to amber. Bank and slip are conventionally displayed on a computer-generatedroll scale and ‘ball-in-tube’ symbol at the bottom of the screen.

74. Since data inputs from systems such as ILS and the radio altimeter are vital to both thedisplayed information and the automatic landing sequence, failure of these data inputs must beannunciated. In EFIS displays this annunciation frequently takes the form of yellow flags painted onthe display screens.

The EHSI Screen75. The lower (EHSI) screen presents a colour display of flight progress in one of nine modes.These are selected on the EHSI section of the EFIS control panel (EHSI Mode Selector Switch) andare MAP, CTR MAP, PLAN, FULL ILS, FULL VOR, FULL NAV, EXPANDED VOR, EXPANDEDILS and EXPANDED NAV.





Map Mode76. This is the display normally used for en-route navigation and is illustrated at Figure 12-12. Itprovides a moving Map display which is orientated to the aircraft’s present track, with the aircraftsymbol positioned at bottom centre and a 60° expanded arc of the compass rose positioned acrossthe top of the screen. Ground features such as navaids, airports and waypoints are shown in theirrelative locations to a common scale (when selected on using the appropriate button(s) on the controlpanel (Figure 12-9). The scale of the Map picture is selected on the EHSI section of the control panel(the Range selector), which typically offers ranges of 10, 20, 40, 80, 160 and 320 nm. The weatherradar picture, generated in the standard colours of green, amber and red (with magenta in somecases), can be superimposed on the display in the EXP VOR, EXP ILS, EXP NAV, CTR MAP andMAP modes, again by pressing the WXR button on the control panel.

77. Heading information is obtained from the aircraft's inertial reference system(s). Whenoperated between the latitudes of 73°N and 65°S the compass rose is referenced to magnetic north ortrue north, depending upon operator preference. Above these latitudes the compass rose is referencedto true north only. Note however that the compass rose lubber line shows aircraft track and that theheading pointer is off centre in conditions other than of zero drift.

78. Wind speed is displayed digitally, with an analogue display of wind direction in the form ofan arrow pointing in the appropriate direction. The wind arrow is oriented to the Map display,which is in turn oriented to the aircraft track such that the vertical axis of the display is the aircraftinstantaneous track as shown at Figure 12-12. The wind velocity shown at Figure 12-12 is thereforein the order of 225°(M)/50 kt.





79. Lateral and vertical deviation from the planned flight path is indicated by pointers and scalesaligned horizontally and vertically on the edges of the display. The selected range scale is overlaidvertically on the Map display, originating from the aircraft symbol. Distance and time to the nextwaypoint is displayed digitally.

80. A trend vector extending from the apex of the aircraft symbol shows the predicted lateralposition at the end of 30, 60 and 90 second intervals based upon bank angle, groundspeed andlateral acceleration. A range to altitude arc intersects the planned track and range scale at the pointwhere a selected target altitude will be reached at present rate of climb or descent.

81. With the EHSI in the map mode the screen is continuously displaying area navigationinformation. The picture is generated by the appropriate signal generator using data provided by theinertial navigation/inertial reference system and by the flight management system. The position of theaircraft as determined by the INS/IRS will be continuously monitored and updated by the FMS usingfixing data received from in-range VOR/DME stations which are automatically selected by the FMS.The FMS selects stations to achieve optimum fix geometry, two DME range arcs at 90° to each otherbeing the ideal. Obviously the automatic update aspect of the area navigation function will ceasewhen the aircraft flies out of VOR/DME coverage. Manual update of the INS/IRS position is possiblebut should not normally be necessary.





FIGURE 12-12Map Mode Display





CTR MAP Mode82. Displays the same data and symbols as the MAP mode, but the aeroplane symbol is placed inthe centre of the map area so that MAP information behind the aeroplane is displayed.

Plan Mode83. Figure 12-13 shows the display generated when plan mode is selected. On the lower part ofthe screen the active route is displayed, but now it is oriented to true north. Track and headinginformation is on an expanded compass rose but now the lubber line shows heading with the trackmark off centre in conditions other than of zero drift. Again distance and time to the next waypointshown digitally. Wind speed and direction is not displayed in this mode and weather radar returnscannot be superimposed. It is a useful display mode for checking route changes as they are selected atthe keyboard and before they are entered into the Flight Management System (FMS) computer.





FIGURE 12-13Plan Mode Display





VOR and ILS Modes

84. These are illustrated at Figure 12-14 and Figure 12-15 respectively. They may be presented asa full compass rose display with conventional heading and course deviation indications, or as anexpanded compass rose display upon which the weather radar picture may be superimposed on a‘semi-map’ picture with the selected range scale displayed. In either case wind speed and directionand system source (ILS or VOR) are annunciated.

85. In expanded VOR and ILS modes (and, indeed, in Map mode) a dotted line appears from theapex of the aircraft symbol to the heading bug for a few seconds following the selection of a new setheading. The aircraft's instantaneous (current) track is displayed as a solid line extending from theapex of the aircraft symbol to the compass scale arc. Bearing of the selected radio navaid is shown bya solid line extending from the centre bar of the lateral deviation scale to the compass arc.





FIGURE 12-14VOR Mode Display





FIGURE 12-15ILS Mode Display

Expanded NAV Mode86. Displays lateral and vertical navigation guidance information similar to a conventional HSI.The FMC is the source of the navigation data. Weather Radar return data is displayed when theWXR Switch is On.





Full NAV Mode87. Displays same data as expanded navigation mode with the following exceptions:Weather Radar displays are not availableA full compass rose is shown in place of the expanded compass rose.

Backup Data Inputs88. In most systems the pilots can, independently of eachother, connect their respective EADI andEHSI displays to alternate sources of input data. For example, should a symbol generator fail on theleft hand side, the captain can duplicate the information shown on the right hand screens. The samecan be done in the event of the failure of either left or right air data computers (ADC) or flightmanagement computers (FMC). In an aircraft equipped with three inertial reference systems (IRS),number one IRS would normally supply the captain’s EFIS and number two IRS the first officer’sEFIS. In the event that either of these IRS were to fail, number three IRS can be selected to replacethe failed system. These selections are made on a source selector switch panel.

EFIS Symbology89. Examples and descriptions of EFIS symbols are given in the tables at Figure 12-16 toFigure 12-25.

90. The following symbols may be displayed on each EHSI depending on EFIS Control Panelselections. General colour presentation is as follows:





GREEN (G) Indicates active or selected mode and/or dynamic conditions

WHITE (W) Indicates present status situation and scales

MAGENTA (M) (pink)

Indications command information, pointers, symbols, and fly-to conditions, weather radar turbulence

CYAN (C) (blue) Indicates non-active and background information

RED (R) Indicates warning

YELLOW (Y) Indicates cautionary information, faults, flags

BLACK (B) Indicates blank areas, off condition





FIGURE 12-16EFIS Symbology





Weather Radar Displays and Annunciations91. Figure 12-12 shows weather radar returns with the EHSI selected to MAP. Figure 12-14 andFigure 12-15 show weather radar returns with the EHSI selected to expanded VOR and ILS modes.

92. The weather radar returns are colour coded red for the most intense returns, yellow for lesserintensity and green for lowest intensity.

93. The radar has three (and possibly four) modes; test, weather (WX), (possibly) weather plusturbulence (WX + T) and MAP.

94. The test mode checks the hardware and paints a predetermined pattern on the screen toassure the operator that the various colours are being properly produced by the EFIS symbolgenerators.

95. The weather mode symbology is as described in the table at Figure 12-25.

96. If incorporated, the weather plus turbulence mode introduces a fourth colour (magenta) ontothe weather paint in areas of suspected high turbulence, which the radar determines by identifyingthe areas of greatest rate of change of target intensity. This is likely to coincide with the area ofgreatest rate of change of vertical velocity of the air, and therefore of greatest turbulence.

97. The map mode employs a vertically broad beam to paint the land/sea surface ahead of theaircraft.

98. When the system is operating normally the radar operating mode (WX, WX + T or MAP, butnot test, as this is self evident) is displayed in the top right corner of the HSI, together with the tiltangle of the scanner. These are not shown at Figure 12-12, Figure 12-14 or Figure 12-15.





99. In the event that anything goes wrong with the weather radar, the pilot is informed by meansof a message which appears towards the bottom left hand corner of the HSI. These messages aretypically as follows, but will vary slightly from one system to another:

VOR/DME Area Navigation (RNAV)

Principle of Operation100. The simplest type of B-RNAV system (used in general aviation) is based on azimuth anddistance information from a VOR/DME. It is also called the RHO-THETA system. With this systemthe pilot effectively moves or off-sets the VOR/DME to any desired location if it is within receptionrange. This ‘phantom station’ is created by setting the distance (RHO) and the bearing (THETA) ofthe waypoint from a convenient VOR/DME in the appropriate windows of the waypoint selector. Aseries of these ‘phantom stations’ or waypoints make up an RNAV route.

WXR FAIL Indicates weather radar has failed (no weather data displayed).

WXR WEAK Indicates weather radar calibration fault.

WXR ATT Indicates loss of attitude input for antenna.

WXR STAB Indicates antenna stabilization is selected off.

WXR DSPY Indicates loss of Display Unit cooling or an overheat condition of the HSI. Weather radar display is blanked.





101. Figure 12-26 illustrates how a VOR/DME RNAV is used to navigate from A to B on a directroute. This route crosses the 180° radial 23 NM south of the ALPHA VOR/DME. Therefore, thepilot sets waypoint 1 as 180/23 on the control panel. Waypoint 2 is 15 NM from BRAVO VOR/DME on the 360° radial, or 360/15 on the panel. Waypoint 3 is 360/22. The direct route from A to Bis 191 NM, 24 NM less than the airways route

FIGURE 12-26VOR/DME RNAV Principle of Operation





102. The pilot could also place waypoint 3 on the destination airport, allowing navigation fromwaypoint 2 direct to B. The DME readout would give a constant indication of the remaining distanceto the destination. The pilot would specify waypoint 3 as 064/49 (based on the CHARLIE VOR/DME). The RNAV computer carries out these vector solutions continuously and displays theappropriate information on the aircraft’s Horizontal Situation Indicator (HSI), Course DeviationIndicator (CDI) and Radio Magnetic Indicator (RMI) such that the aircraft can be flown to anywaypoint along a direct track.

103. Simple B-RNAV equipments use the input from one VOR/DME at a time, whereas moresophisticated B-RNAV systems use two or more VOR/DME stations for more accurate positionresolution. You will recall that modern Flight Management Systems (FMS) are programmed with thelocation of each VOR/DME (and ILS/DME) and are capable of automatic selection of the mostsuitable beacons for any planned route fed into the FMS computer programme.

Advantages104. The main advantage of the VOR/DME (B-RNAV) system is that it enables the pilot to flydirect to a given location, or a series of locations, using ground stations which are not situated atthose locations. Full use of the available airspace can therefore be made subject to the availability ofphantom way-points. Please note that phantom stations can only be defined within the range of theVOR/DME stations adjacent to the RNAV route.





Errors105. The use of a VOR in a B-RNAV solution outside of the designated operational coverage(DOC) can lead to serious navigational errors. This is particularly to be noted when using more thanone VOR/DME in the B-RNAV configuration, when it is virtually impossible for the pilot topositively identify which beacon is currently being used by the B-RNAV equipment, and therefore toestablish whether or not the information being produced is reliable.

106. Accuracy is, of course, no better than the accuracy of VOR (± 5°) and DME (± ¼ nm plus1.25% of slant range). Close to the DME ground transponder the error due to slant range will begreatest, similarly any deflection of a VOR radial will have an adverse effect on the RNAV computedposition.





B-RNAV Control and Display UnitFIGURE 12-27Typical RNAV control and display unit

107. Figure 12-27 shows an example of a B-RNAV control and display unit (CDU). The controlsconsist of a Mode Selector Switch, a Display Selector Switch and a Keyboard.

108. The Mode Selector determines in which of three modes the equipment operates the aircraft’sHSI steering commands.





VOR/LOC. Conventional navigation in which the ground station is the waypoint and bearingto/from is a VOR radial.

RNAV. The waypoints are not at the ground station position. The computer generates rangeand bearing from an offset VOR/DME. Within 100 nm range the HSI deviation indication is linearwith full-scale deflection at 5 nm left or right of track (i.e. 2nm/dot). Beyond 100 nm the deviationindication is angular (degrees left or right of track).

APR (Approach). The mode of operation is the same as for RNAV, but with linear deviation upto 25 nm range and full-scale deflection at 1.25 nm left or right of track. This mode will provideenhanced steering command accuracy for approaches to a non-beaconed location.

The TEST function produces a specific display.

The Display Selector controls the numerical displays of the CDU.

SBY. Standby waypoint information in terms of station frequency, bearing and range from thepreceding waypoint, station elevation x 100 ft and course is displayed.

ACT. Active (in-use) waypoint information is displayed to the same parameters as SBY.

BRG/DST. Displays bearing and range to the in-use waypoint in RNAV and APR mode bysolving first the slant range triangle (hence the need for station elevation) and then the RNAVtriangle. In VOR/LOC mode bearing and range displayed is to the VOR/DME station.





KTS/TTS. Displays groundspeed (in BRG/KTS window) and time-to-station in minutes (in DST/TTS window). In RNAV and APR the time is to waypoint; in VOR/LOC it is to VOR/DME groundlocation.

The course display may be either inbound or outbound; either IN or OUT will be illuminated next tothe CRS window as appropriate.

Data may be entered into the computer by means of the keyboard or by prepared magnetic card, anda card reader.

Mandatory Carriage of RNAV Equipment109. What follows is a summary of the relevant parts of UK AIC 148/1997 (Yellow 280).

110. RNAV has been identified as the future navigation system in the ICAO European region.This will in the future mean that routes (airways, upper ATS routes, advisory routes and arrival anddeparture routes) will not necessarily be constrained to run between point source navigation aids(principally VOR/DMEs). Straightening airways and other routes (by making it unecessary to flybetween one VOR/DME and the next) will result in considerable fuel savings for operators.

111. It should be noted that, currently, the primary source of the position data for B-RNAV isVOR/DME. Alternatives, including GPS, have been accepted by the Joint Aviation Authorities (JAA)as a source of B-RNAV position data.

112. Many aircraft operating within European airspace are required to carry suitable RNAV

equipment (basic or otherwise) with effect from 23rd April 1998. Note that the requirement to carryVOR, DME and ADF equipments remains unchanged.





B-RNAV Implementation - ECAC Airspace113. With effect from 23 April 1998, the carriage of B-RNAV equipment approved for RNP 5operations will become mandatory on the entire ATS Route Network in the ECAC area includingdesignated feeder routes (SIDs and STARs) in/out of notified TMAs.

114. As of 23 April 1998, aircraft, other than State aircraft operating on the ATS Routes, abovethe lowest applicable flight level as published by States, shall be equipped with, as a minimum,RNAV equipment meeting RNP 5 in accordance with the requirements set out in ICAO Doc 7030Regional Supplementary Procedures (EUR. RAC section 15).

Note: The lowest applicable flight level might vary throughout the 36 ECAC States, but none areknown to be mandating B-RNAV below FL 100, and many will accord with the UK.

115. As of 23 April 1998 no exemptions, other than for contingency situations, will be given,regardless of whether such exemptions were offered in earlier AICs by some States.

116. Operators of aircraft fitted with RNAV equipment having a navigation accuracy meetingRNP 5 shall insert the designator letter "R" in Item 10 of the Flight Plan.

117. Having the capability to operate on RNP 5 routes defined by VOR/DME does not imply thatthe aircraft is suitably equipped to operate on B-RNAV routes in the ECAC area.

118. The Eurocontrol RNAV Standard Doc 003-93, Area Navigation Equipment OperationalRequirements and Functional Requirements (RNAV), defines the functional requirements for B-RNAV equipment.





B-RNAV in UK Airspace119. Within UK Airspace, for an as yet unspecified period of time but in accordance with thedecision of many of our European neighbours, B-RNAV will be implemented only above FL 95 (thelowest FL at which carriage of B-RNAV equipment will be mandatory is FL 100).

120. Moreover, for an as yet unspecified period, mandatory B-RNAV procedures will not beapplied to any designated feeder routes (SIDs and STARs) in/out of UK TMAs.

121. These indeterminate times will be reviewed periodically in the light of the future ATS Route splanning.

UK ATS Routes122. Historically ATS Routes have been delineated by ground based navigation aids (todaypredominately by VORs). From 23 April 1998 all ATS Routes in the UK will be defined by WGS 84geographical points which may not be coincident with a VOR. This has the advantage that if a VORis moved or withdrawn, the alignment of the ATS Route can remain unchanged. No immediaterealignments of ATS Routes are envisaged due to the mandate of B-RNAV.

123. In the introduction to the ATS Route Catalogue published in the new UK AIP, a statement ofthe Required Navigation Performance (RNP) will be given, together with an explanation as to howthis will be applied. Although in general UK ATS Routes will be available for non RNAV equippedaircraft operating below FL 100, certain UK ATS Routes will be RNAV Routes at all flight levels. Amethod of readily identifying such routes will be published. Where appropriate, notes relating toindividual ATS Routes will be amended.





124. Route designators are allocated according to Annex 11 to the Convention on InternationalCivil Aviation. The allocation of international route designators in ECAC States is arranged throughICAO EUR (Paris) and co-ordinated by the Route Network Development Sub-Group (RNDSG) ofEurocontrol. There are currently insufficient RNAV designators to retitle all RNAV routes andarrangements are being co-ordinated through ICAO to resolve this shortfall.

125. For practical operational reasons, and on safety grounds, it is inappropriate to change alldesignators at the same time. Therefore, the re-designation of routes will be introduced in phases.As a consequence, in the medium term, a mixture of conventional and RNAV designators will beused in the UK even though all ATS Routes will be RNAV above FL 95. Precedence for change willbe given to those ATS Routes whose alignment is changed and to those requiring international co-ordination.

Certification and Approval Requirements126. To be eligible for B-RNAV operations, on-board navigation equipment will be required toprovide en-route lateral track keeping accuracy of +/- 5 nm or better for 95% of the flight time (RNP5).

127. The JAA has published TGL No 2, rev 1, giving certification and approval guidelines for B-RNAV installation. This TGL is acceptable to the UK CAA and may be used as a certification basis.

128. For UK operators and UK registered aircraft the only approval required under the ANO isthat B-RNAV equipment and its installation in the aircraft have been approved in respect of anyaircraft or specified class or category of aircraft or in respect of a specified type or types of equipment(ANO Article 43). No separate operational approvals are required.





Provision of Navigational Infrastructure129. Provision of a navigation infrastructure, which will enable users to achieve the requirednavigation accuracy, will remain the responsibility of States.

130. Until at least 2005, ECAC member States will continue to provide the VOR/DMEinfrastructure necessary to enable operators to meet the required system use accuracy. Operatorsshould be aware that DME is expected to become the primary source of position information in theECAC area and the maintenance of VOR beyond 2005 may not be guaranteed (ICAOImplementation Strategy of the Future Air Traffic Management Systems in the European Region(FEATS), Part 2, paragraph 2.2.2 refers).

Responsibility of Operators131. The navigation system accuracy achievable by an RNAV system is dependent upon both theairspace infrastructure and the airborne equipment. It is the responsibility of the operator to ensurethat the required system accuracy can be achieved when planning to operate in designated B-RNAVAirspace.

Note: Where position derived from GPS is the only input to the RNAV system it is incumbent uponoperators to confirm that the necessary coverage from GPS is provided for the intended flight (JAATGL No 2 rev 1 paragraph 5.2 refers).





NOTE:

(THIS IS THE END OF THE RADIO NAVIGATION SYLLABUS)





Doppler

Doppler Effect

Doppler Frequency Shift Calculations

Airborne Doppler Systems

Single-beam Systems

Two-beam Systems

Three-beam systems

Four-beam Janus System

Features of the Janus System

Beam Shape

Doppler Errors

Accuracy

Airborne Equipment

Other Doppler Applications



Doppler


13Doppler

1. The Doppler principle is used to establish the relative speed of a moving object bymeasurement of a received radio frequency and the determination of the difference between thetransmitted and the received frequencies. Practical examples of its use are Moving Target Indication(MTI) systems in surveillance radars, Doppler VOR and the measurement of an aircraft’s drift andgroundspeed.

Doppler Effect2. Doppler effect takes its name from the nineteenth century Austrian physicist ChristianDoppler, who predicted it in connection with light waves. It is based upon the principle that areceived frequency will only be the same as the transmitted frequency provided there is no relativemovement between receiver and transmitter. Let us consider a simple example.

3. Suppose you are at the seaside standing knee-deep in the sea. The waves are rolling in at therate of one every five seconds – so you will be receiving a cold slap in the belly at that frequency –once every five seconds.

4. If you now walk forward into the sea the waves will strike you with increased frequencybecause the relative velocity between you (the receiver) and the sea (the transmitter) is positive – ie:you are moving towards each other. The faster you move, the greater the frequency with which thewaves will strike you, although their rate of transmission (transmission frequency) hasn’t changed.Thus, it can be seen that the difference between transmitted and received frequencies is directlyproportional to the relative velocity between receiver and transmitter.



Doppler


5. If you were to walk backwards out of the sea, the frequency with which the waves struck youwould be less than once every five seconds – ie: lower than the transmitted frequency, becauserelative velocity is now negative.

6. Exactly the same principle applies to soundwaves (sonic frequencies), radio waves (radiofrequencies) and, for that matter, light waves (above radio frequency). If a train, sounding itswhistle, is moving towards an observer, the pitch of the whistle tone sounds high to the observerbecause there is positive relative velocity between the two, so the received frequency is higher thanthe transmitted frequency. As the train passes the observer the apparent pitch of the whistle tonefalls sharply, as the relative velocity becomes first zero and then negative.

7. The difference between the received frequency and the transmitted frequency is known as theDoppler Shift, or beat frequency.



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FIGURE 13-1The Doppler Principle

8. At Figure 13-1 the principle is considered diagramatically. Transmitter A is moving inrelation to two stationary receivers, B and C. The distance between the wave fronts passing receiverC is decreased, therefore since effective wavelength has decreased, apparent (received) frequencymust have been increased. Conversely, the received frequency at B will be lower than the transmittedat A.



Doppler


Doppler Frequency Shift Calculations9. The difference between the transmitted frequency and the received frequency is known as theDoppler frequency shift and is directly proportional to the relative velocity between transmitter andreceiver. If the range between the transmitter and receiver is decreasing, the Doppler shift will bepositive (the frequency received will be higher than the frequency transmitted). Conversely if therange between the transmitter and receiver is increasing, the Doppler shift will be negative.

10. The formula for calculating Doppler shift is a simple one. The formula states that:

Where

Ds = Doppler shift (Hz)

S = The relative speed (metres/second) between transmitter and receiver

λTX = The wavelength of the transmitted signal (metres)

DsS

λTX----------=



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EXAMPLE 13-1EXAMPLE

A stationary transmitter is producing a signal at frequency of 5 GHz. A receiver ismoving directly towards the transmitter at 600 km/hr. Determine the Doppler shift.

SOLUTION

=

=

=

=

=

Ds=

=

S (metres / second) 600 km/hr 1000×60 60×

-------------------------------------------

167 metres/second

λtx(metres) CF----

38×10

59×10

--------------

0.06 m (or 6 cm)

167 m/sec0.06 m

------------------------

2783 Hz and the shift is positive



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EXAMPLE 13-2EXAMPLE

A transmitter is moving straight towards a stationary receiver and is transmitting at a frequencyof 10 GHz. The measured Doppler shift at the receiver is +5 KHz. Determine the speed in knotsof the transmitter.

SOLUTION

= CF

=

= 0.03 m (or 3 cm)

S =

= (5 x 103) x 0.03 m

=

= 150 x 1.95 kt

= 292 ½ kt

λTx (metres)

38×10

109×10

-----------------

Ds λTx×

150 m/sec



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Airborne Doppler Systems11. In an airborne Doppler system the transmitted energy must be directed downward to strikethe ground beneath the aircraft as in Figure 13-2.

FIGURE 13-2Beam Depression in a Typical Doppler System

12. The amount by which the beam is depressed is inevitably a compromise; a large angle ensuresadequate returned signal but a low scaling factor (ie doppler shift per knot very low), whereas a smallangle gives a better scaling factor but less returned signal. Values chosen practically vary between60° and 70°. (see Figure 13-2).



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Choice of Frequency13. As the transmitted frequency is increased, the doppler shift for a particular groundspeedincreases. A rule of thumb is that the doppler shift is 34 Hz per 100 MHz of transmitted frequencyper 100 kts groundspeed, multiplied by the cosine of the depression angle for a single beam system asin Figure 13-2. Two bands of frequencies are allocated for airborne doppler use, 8,750-8,850 MHzand 13,250-13,400 MHz.

Beam ArrangementsFIGURE 13-3Principle of Operation of a Single-Beam Doppler System



Doppler


Single-beam Systems14. The finite horizontal beam width of the emitted beam causes not one discreet dopplerfrequency, but a whole spectrum to be observed at the receiver. With reference to Figure 13-3, thecentre of the beam AB is pointing along track and the doppler shift reduces towards the edges of thebeam (at points A and B the doppler shifts are equal).

15. The shift observed from beam CD, pointing off track by δ, varies throughout the beam width.

16. Hence the nature of the returned signal will change if the beam is swept in azimuth. Themean doppler shift is maximum and the width of the frequency spectrum minimum if the beam ispointing along the track. As the beam is turned off track the mean doppler shift decreases and thespectrum broadens.

17. It is possible to utilise these phenomena in a simple single-beam doppler system. The beamcould be rotated until Ds was maximum, drift noted and groundspeed computed, the parameters ofthe radar being known. This would be a crude and inaccurate system and is not used in practice, butthe property of the changing width of frequency spectrum can be utilised in a practical drift-sensor.



Doppler


Two-beam SystemsFIGURE 13-4Fixed, Two-Beam Doppler System



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FIGURE 13-5Moving, Two-Beam Doppler System

18. The doppler shifts observed in two beams can be combined to give an automatic derivation ofdrift and groundspeed. The beams may be fixed to the aircraft axes as in Figure 13-4 or rotated inazimuth as in Figure 13-5; the beams also could be fixed but have the same general layout as inFigure 13-5.

19. The two-beam Doppler system, works well in hovercraft (Marconi and Ryan both produceworking systems). Here the two fixed beams are directed into the wake that exists behind the craft.The beams are depressed by a shallow angle of about 45° to give a high scaling factor and make thesystem accuracy comparatively unsusceptible to pitch changes.



Doppler


Three-beam systems20. A third beam could be directed, say, vertically down to sense and eliminate errors due tovertical motion, but such systems are not generally used. Many modern dopplers are in fact three-beam systems, but take the form of a Janus four-beam system (see below) in which one beam, beingredundant is removed.

Four-beam Janus System21. The four-beam system, known as the Janus system after the Roman god portrayed with twofaces looking in opposite directions, has beams looking both forward and backward. The aerial maybe rotatable in azimuth or fixed.



Doppler


FIGURE 13-6Moving Janus Beam Layout



Doppler


Moving Aerial22. The general layout of a moving Janus system is shown in Figure 13-6. The frequency receivedfrom beam A is mixed with that received from C. The frequency observed from A will be higher thanthe transmitted frequency by Ds (A) and that received from C will be lower than the transmittedfrequency by the same amount. Hence the total doppler shift, Ds (A – C), will be twice thatobserved in either A or C separately. The aerial is initially rotated in azimuth until Ds (A-C) = Ds (B-D); it is then aligned with track and the angle between the fore and aft axis of the aircraft and theaerial will represent drift.



Doppler


FIGURE 13-7Fixed 3-Beam Janus Layout



Doppler


Fixed Aerial 23. Some lightweight doppler systems have a fixed Janus aerial employing only three beams, asdepicted in Figure 13-7. If the doppler shift is derived individually from each beam it is possible todetermine aircraft velocity along all three axes by the subsequent doppler mixing.

Fixed v’s Moving Aerial 24. Similar information is available from moving four-beam and fixed three-beam systems, butfor a compact and comparatively reliable and robust system with an almost-instantaneous responseto aircraft flight path changes (essential for a helicopter doppler) the fixed aerial is far more suitablethan the moving type. A moving aerial may give better accuracy, however, as mentioned later.

Features of the Janus System25. Almost all airborne dopplers use Janus aerials. The advantages of this system over a single-ended system are discussed below.

Doppler Frequency Measurement 26. The doppler frequency shift from a Janus aerial is twice that from a single-ended system.Hence scaling-factors (Hz per knot) can be made correspondingly larger, with an increase in possibleaccuracy of measurement of doppler frequency and hence speed.



Doppler


Transmitter Instability 27. It can be proven that, within certain limits, the doppler shift as seen by a Janus array isindependent of transmitter frequency (ie: transmitter frequency stability is not so much of a problemin this equipment).

Vertical Motion 28. Any vertical motion being experienced by an aircraft will affect both the forward andrearward beams of a Janus array by the same amount and in the same sense. When the twofrequencies are subtracted from each other any component of doppler shift due to the vertical motionof the aircraft is cancelled out.



Doppler


FIGURE 13-8Effect of Pitch Changes on a Janus Array



Doppler


Pitch Changes 29. With reference to Figure 13-8 it can be shown that any pitch change causes a greaterdepression of one beam, with an associated reduction of Ds, but reduces the depression of the other,increasing Ds in that beam. This causes very little total error in measured groundspeed.

Beam Shape30. The description of the doppler beams so far has assumed pencil beams depressed by σ in avertical plane and displaced by φ= in a horizontal plane, both with respect to the aerial axis. Inpractice the beams are produced by linear arrays, and are conical about the waveguide axis. Thesemi-angle of the cone is defined as the ‘depression angle’, and usually denoted by θ.



Doppler


FIGURE 13-9Practical Doppler Beam Shape



Doppler


FIGURE 13-10Pattern of ‘Isodops’ produced by Doppler

31. Naturally only parts of the conical beam are used, as shown in Figure 13-9, but at what everrange and angle these beams strike that surface beneath the aircraft, the doppler shift will always beproportional to cos θ. Where the cone cuts a plain surface beneath the aircraft, a hyperbola of equaldoppler shift, or ‘isodop’ is formed, as shown in Figure 13-10. The beam width along the isodop willnot affect the doppler shift, and the aerial may be rolled with no affect on the doppler frequency; thebeams simply roll along the isodops. Hence very rarely are doppler aerials roll-stabilised, althoughthere is obviously a limit to the amount of roll with which a system can cope due to lifting of theuppermost beam. It is not always correct to say that roll does not affect the doppler shift; if there isdrift and roll and/or pitch present the beams are rolled about the aircraft axis and not the aerial axisand do not follow the isodops, causing a small error. This is discussed under ‘Errors’.



Doppler


Doppler Errors32. Having mentioned the principles of doppler it is now possible to consider how errors may beintroduced into the final output. The main sources of error are discussed below.



Doppler


Sea or Terrain BiasFIGURE 13-11Response Curves for Varying Depression Angles and Surfaces



Doppler


FIGURE 13-12Received Frequency Spectrum - Land vs Water



Doppler


33. It is normally assumed that the amplitude of the doppler frequency spectrum is symmetricalabout its central frequency. With reference to Figure 13-11 it can be seen that over land thisassumption is reasonable. The figure indicates that over land the radar cross section per unit areawill be virtually constant over a typical 4° depression beam-width centered between 60° and 70°.Over sea, however, the curves have a pronounced slope and even over a fairly rough sea (B4) a beamdepressed by, say, 67° would have quite different amplitudes returned from points ± 65° and 69°. Thespectrum will now be unsymmetrical as indicated in Figure 13-12 and any form or tracker will tracka frequency lower than that tracked over land.

FIGURE 13-13Sea Bias Error Values

34. The sea bias error, values of which are shown in Figure 13-13, is proportional to the squareof the depression beam width. To minimise the error one must have either a very narrow depressionbeam-width (eg Marconi AD 560, AD 570) or incorporate a Land/Sea switch that alters the effectivescaling factor to suit conditions (eg Decca 62M, 67M).

Depression Beamwidth

% Error Under Read

Beaufort 2 Beaufort 4

2° 0.26 0.18

3° 0.56 0.41

4° 1.00 0.72

5° 1.58 1.13

6° 2.30 1.63



Doppler


Sea Movement Error35. Dopplers measure velocity relative to the terrain below the aircraft. If the sea beneath theaircraft is moving, its motion must be considered when deriving position from the dopplerinformation. Sea movement takes two basic forms mentioned below:

Tidal Flow36. Tidal flow is normally less than 2 kt. It may be corrected-for, with knowledge of localsituations, by applying a down-stream vector to the indicated doppler position. Tidal streams mayreach up to 10 kt in narrow channels but these conditions are unlikely to affect an aircraft for anyappreciable time.



Doppler


FIGURE 13-14Typical Drift Correction Graphs



Doppler


Water Transport37. Although wave motion basically involves vertical motion only, wind across the surface ofwater causes eddies with motion downwind on the surface and upwind below the surface. Thedoppler energy is reflected by the surface, and hence measurements are made relative to this movingsurface. Correction can be made by correcting the ground position with a downwind vector usingsurface wind direction and one-fifth surface wind amplitude; maximum amplitude should represent a6 kt correction (eg for 30 kts and above). Both groundspeed and drift are likely to be in error; anexample of a drift correction graph is shown in Figure 13-14. Some equipments have facilities forsetting in surface wind corrections and hence indicating corrected ground position.

Flight Path and Pitch Change Errors38. Errors due to combinations of flight path and pitch changes have been covered previously. Inshort, the errors are very small if a Janus aerial is used, hence many systems have aerials unstabilisedin pitch. Gyro stabilisation should completely eliminate errors, flight path stabilisation shouldalmost eliminate them; data stabilisation implies compensating for errors in the computer acceptingthe raw doppler output.



Doppler


Roll Error39. Roll normally has little effect on doppler accuracy. With a moving aerial the effect of roll isto move the beams along the isodops, causing no error. A very small error in drift will resulthowever, if much drift is present; the isodops are symmetrical about aircraft track but the beams rollabout aircraft heading. This situation applies equally to fixed aerial systems. Roll can have a lot ofeffect on the ability of a doppler to remain locked-on to returned signals; this is due to lifting of theuppermost beams to produce large incidence angles. The most critical situation is naturally at highlevel over a calm sea. Operational limitations are often stated as maximum roll angles Vs height. Eg± 30° at 20,000 ft and ± 20° at 40,000 ft (Decca 62 M).

Drift Error40. The presence of large drift angles will have no detrimental effect on the accuracy of a movingaerial system within its limits.

Height Error41. Height errors exist due to the fact that the doppler measures spacial velocity (if at height) andnot the velocity with which the ground passes at a point immediately beneath the aircraft.

42. The groundspeed and distance indicators will overread at height with respect to the rate atwhich the surface is moving horizontally at a point vertically below the aircraft. The indicators do,however, give correct spatial velocity. The error is small, about 0.2% at 40,000 ft, but may well belarger than any other inaccuracies in a modern doppler. It should be compensated for on a long legat high level. Values for the error are given in Figure 13-15.



Doppler


FIGURE 13-15Height Error Correction Figures

Height-Hole Error43. The discussion so far has centred around pulsed Doppler Systems however a CW system is themost efficient possible as regards utilisation of transmitted power. This system requires separatetransmit and receive aerials, however, and in the past considerable difficulty has arisen from thepresence of unwanted cross-coupling between transmitter and receiver. Frequency modulation hasbeen used as solution to the unwanted cross-coupling that can occur in a pure CW system.Theoretically duplex operation is possible with a single aerial, but practically this has not beenachieved.



Doppler


44. If the round-trip path-length travelled by an FMCW transmission is an integral number ofmodulating wavelengths there will be no output when the received signal is mixed with a sample ofthe transmitted signal. At these points, ‘height-holes’ or ‘altitude-gaps’, the doppler will cease tofunction.

45. The height-hole problem may be overcome in two ways :

(a) Wobbulation If the modulation wavelength is continuously changed an aircraft willnot remain in a condition of no signal for any appreciable time. The process ofcontinuously changing the modulating frequency (fm) is ‘wobbulation’. As anexample, the Marconi AD 2300 has an fm of 200 kHz wobbulated by ± 16% at 10cycles per minute.

(b) Broadside Beamwidth If the broadside beam width of a doppler is made large, aheight-hole occurring in one part of the beam will tend to be masked by returnscoming from elsewhere in the beam at different slant ranges.

Other Errors46. Variation in waveguide dimensions is a source of error not mentioned; current engineeringpractice, however, gives errors in cos θ of a small fraction of 0.1%. Accurate aerial alignment inazimuth is important if systematic errors are to be avoided. Alignment is normally better than 0.1°.

Accuracy47. Current Doppler equipments give accuracy figures in the order <0.1% in distance gone and<0.2° in drift ; in other words, the along track error is much smaller than the across track error andthis will result in an elliptical shape of position errors.



Doppler


Airborne Equipment48. Figure 13-16 and Figure 13-17 show a typical Doppler Aerial installation and flight deckdisplay as used on many older aircraft types. The associated control panel (not shown) is very simpleto operate and has 3 separate switches, the functions of which are described below:

(i) Power ‘OFF/STBY/ON’. Selecting the power to ‘STBY’ allows the equipmentto warm up to normal operating temperature whilst on the ground, but doesnot allow the TX to transmit until the ‘power on’ selection is made, shortlyafter the aircraft has passed VR.

(ii) 4 position ‘slew’ switch. Two of the positions on this switch are used formoving (inching) the drift to its calculated value at the point where the TX isswitched on ; the other two positions provide the same function in terms ofgroundspeed. In other words the slew switch is used to inch the drift andgroundspeed values to approximately the correct values on take-off such thatthe equipment can ‘lock-on’ to the appropriate signal and hopefully thereaftertrack both parameters automatically. In the event of insufficient signal beingreceived during flight, the slew switch can be used to set DR values of drift andgroundspeed on the display until such time as the signal is regained.

(iii) Land/Sea switch. The purpose of the Land/Sea switch is described inparagraph 34.

49. The student should note that there are two flags that could appear on the display unit shownin Figure 13-17; a yellow / black striped flag appears if the power supply to the equipment fails, anda white flag with the letter ‘M’ will appear whenever the received signal strength has fallen below apre-set level and the equipment has reverted to ‘memory’ mode (i.e. frozen display values).



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FIGURE 13-16Typical Aerial Installation



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FIGURE 13-17Typical Flight-Deck Display

Other Doppler Applications50. Ground radars make use of the Doppler principle to eliminate radar returns from fixedobjects such as hills, buildings, masts and so on. The process is known as moving target indication(MTI). The principle is that returns from moving targets suffer a doppler shift whereas returns fromstationary targets do not.

51. Ground DF stations can also employ Doppler techniques in order to improve accuracy andminimise siting errors.



Doppler


52. Doppler VORs are also less susceptible to siting errors. Doppler VOR was discussed in thechapter on VOR.



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On a typical ATC PPI display, aircraft targets are shown as a blip on the screen. However, with amore modern display this blip may be replaced with a single letter - what does the letter indicate:

QUESTION 2.

What frequency does the airborne SSR equipment transmit on:

QUESTION 3.

To facilitate its main function an SSR transmits a series of pulse pairs however, a third pulse is alsousually transmitted. What is the purpose of this third pulse:

QUESTION 4.

Excluding the SPI pulse, how many pulses go to make up the total air-to-ground SSR pulse profile:

QUESTION 5.

How long does the SPI pulse remain once the SSR IDENT button has been depressed:

QUESTION 6.

What does code A7007 signify when selected on an SSR:



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QUESTION 7.

What are the increments in height when using Mode C:

QUESTION 8.

What would the figures "18Z " mean if displayed in the height block on an air traffic controllersradar screen:

QUESTION 9.

Name the two types of interference which can be experienced by Mode A/C SSR systems:

QUESTION 10.

What information does a Mode S "All Call " response contain:

QUESTION 11.

What information is reported automatically under the Basic Functionality of Mode S:

QUESTION 12.

What does the term "squitter" mean in relation to Mode S:

QUESTION 13.

Doppler operates on the principle that __________________ between a transmitter and receiver willcause the received frequency to ____________ if the transmitter and receiver are moving____________.



Doppler


QUESTION 14.

A stationary radio transmitter is transmitting a 5 GHz signal. A receiver is moving directly towardsthe transmitter at 800 km/hr. The doppler shift measured at the receiver will be:

QUESTION 15.

What is a typical beam depression angle in a Doppler equipment:

QUESTION 16.

What is the advantage, in terms of groundspeed measurement, of a Janus aerial over a single beamsystem:

QUESTION 17.

A flight over calm sea would probably produce an _____________ in calculated groundspeed:

QUESTION 18.

What is the usual method of compensating for sea bias error in a Doppler equipment:

QUESTION 19.

Name the two types of Sea Movement Error that may be experienced by a Doppler equipment:

QUESTION 20.

Why does a roll manoeuvre have little effect on the accuracy of a Doppler equipment fitted with aJanus aerial:



Doppler


QUESTION 21.

Does Height Error in a Doppler system result in an overread or an underread in the groundspeed anddistance indicators:

QUESTION 22.

Name one of the two methods that are used to eliminate the height hole problem in an FMCWDoppler equipment:

QUESTION 23.

What are the accuracy figures for distance gone and drift in a modern Doppler system?

QUESTION 24.

What is the purpose of the "slew" switch on a Doppler control panel:

QUESTION 25.

When will the "Memory Flag" be displayed on a Doppler drift and groundspeed display:

ANSWERS:ANSWER 1.

The letter indicates which controller is assigned to that particular target at that time

ANSWER 2.

1090Mhz



Doppler


ANSWER 3.

The Sidelobe Suppression Control Pulse (SLS) is used to suppress replies to any sidelobeinterrogations of an SSR

ANSWER 4.

14 pulses - two frame pulses plus 12 selectable pulses

ANSWER 5.

20 seconds

ANSWER 6.

A7007 signifies that the aircraft is being operated under the Open Skies Treaty arrangements

ANSWER 7.

Mode C increments every 100 ft

ANSWER 8.

The aircraft is flying at 1800 ft relative to the QNH

ANSWER 9.

Fruiting and Garbling

ANSWER 10.

Aircraft Identity plus the capability of the onboard equipment



Doppler


ANSWER 11.

Aircraft Identity (callsign), transponder capability, altitude reporting (25 ft increments), flight status(airborne or on the ground)

ANSWER 12.

Squittering is when an airborne SSR transmitter automatically transmits pre-formatted informationwhich is not in response to an interrogation

ANSWER 13.

relative motion decrease apart

Relative motion/decrease/apart (or, Relative motion/increase/towards each other)

ANSWER 14.

= = 0.06 m = 222 m/sec

= 3700 Hz

3.7 MHz

ANSWER 15.

Beam depression angles are usually between 60° and 70°

DsS

λTX----------= λTX

3 108×5 109 8( )×------------------------= S

800 1000×60 60×---------------------------=

2220.06----------



Doppler


ANSWER 16.

We get twice the Doppler shift from a Janus system and can therefore evaluate groundspeed moreaccurately

ANSWER 17.

A calm sea would result in a lower groundspeed being calculated and displayed since some of theenergy would be reflected in other directions rather than returning to the aerial

ANSWER 18.

A "Land/Sea" switch is used to compensate for sea bias error

ANSWER 19.

Tidal Flow and Water Transport error

ANSWER 20.

The effect of roll is to move the beams along the ISODOPS thereby producing no error

ANSWER 21.

Overread (approximately 0.2% at 40,000 ft)

ANSWER 22.

WOBBULATION (variation of the modulating frequency) or the use of a LARGE BROADSIDEBEAMWIDTH



Doppler


ANSWER 23.

Distance gone < 0.1%

Drift < 0.2°

ANSWER 24.

The "slew" switch is used to set DR values of drift and groundspeed if the reflections from the earth’ssurface are not strong enough to permit the equipment to "lock-on"

ANSWER 25.

If the returning signal strength falls below a preset level the equipment will revert to "memory" andthe drift and groundspeed displays will be frozen at the latest computed values.






Hyperbola

Accuracy in a Hyperbolic Lattice

Range in a Hyperbolic Lattice





14Hyperbolic Navigation System Theory

1. Hyperbolic navigation systems cover wide areas and basically consist of a ‘grid’ of positionlines superimposed upon a geographical area; position is established with reference to the grid lines.

2. We are all familiar with the numbered grid lattice superimposed upon an Ordnance Surveymap, by which position can be quantified in terms of Eastings and Northings. A hyperbolic grid isbasically the same, except that the grid reference lines are not all straight and they do not necessarilyintersect each other at right angles.

3. First, let us consider how hyperbolae are constructed geometrically and then see how this isuseful in determining geographical position using radio signals from fixed ground locations.

Hyperbola4. A hyperbola is a line joining points of equal difference of distance between two fixed points.

5. To fully understand the above definition it is perhaps easiest to construct a hyperbolic lattice.

6. Refer to Figure 14-1 which shows two fixed points (A & B) 100 nm apart. The firsthyperbola is already constructed and is labelled ‘0 nm’. This particular hyperbola is the rightbisector of the base line joining A and B. At any point along the right bisector the distance betweenthe point in question is the same to A as to B, or in other words, the difference of distance is zero.





FIGURE 14-1Basic Geometry in Hyperbolic Navigation Systems

7. Figure 14-2 shows the next hyperbola in the family. In this case any point along the dottedhyperbola is 20 nm closer to A than to B, hence the designation +20 nm. Notice that, at the baseline, this hyperbola is 10 nm from the right bisector.





FIGURE 14-2Construction of Hyperbola

8. Figure 14-3 shows the mirror image of the previous hyperbola, but in this case all pointsalong the dotted hyperbola are 20 nm closer to B than to A, hence the designation -20 nm.





FIGURE 14-3Mirror Image of a Hyperbola

9. Figure 14-4 shows the complete family of hyperbola drawn at every 20 nm difference ofdistance. Notice that the base line extensions (the base line extended beyond A and B) are themselveshyperbola since any point along either base line extension is, in this case, 100 nm closer to one of thetwo fixed points (A or B) than to the other.





FIGURE 14-4Hyperbolic Lattice Construction Between Two Stations A+B

10. Questions may be set involving a simple calculation concerning hyperbolic lattices. Anexample follows:

Example 111. An aircraft is situated on a hyperbola constructed from two stations 120 nm apart. Theaircraft is 70 nm from station A and 90 nm from station B. How far is the hyperbola passingthrough the aircraft from station A at the point where the hyperbola crosses the base line?

Solution12. The situation is as depicted at Figure 14-5.





FIGURE 14-5Solution to Example 1

13. The right bisector is shown in Figure 14-5 since it makes the solution simple. The aircraft is70 nm from station A and 90 nm from station B. The value of the hyperbola is therefore +20 nm. Inother words at any point along the hyperbola the distance to A is 20 nm less than the distance to B.Remember that the distance along the base line from the right bisector to any hyperbola is alwayshalf the value of the hyperbola in question. In this example the distance along the base line from theright bisector to the + 20 nm hyperbola is therefore 10 nm and the distance from station A to thehyperbola is 50 nm along the base line.

Accuracy in a Hyperbolic Lattice14. Figure 14-6 shows a hyperbolic lattice in which it is assumed that the error in the finalposition line may be as much as 10% of the distance between the individual hyperbola. Check thediagram and satisfy yourself that the following statements are correct:





On any hyperbolic lattice:

(a) The system is most accurate along the base line.

(b) If you depart from the base line the area surrounding the right bisector givesreasonable accuracy, and the error in distance for a given percentage error increaseswith departure from the right bisector towards the base line extensions.

(c) The error increases with range from the base line.

(d) It is inadvisable to use the hyperbola adjacent to the base line extensions since a risk ofambiguity of position exists, see Figure 14-7. Here the aircraft's position isestablished using the 180° radial from VOR ABC and the + 90 nm hyperbola. It isimpossible to say whether the aircraft is in fact at position X or Y.





FIGURE 14-6Accuracy within a Hyperbolic Lattice





FIGURE 14-7Area of Ambiguity within a Hyperbolic Lattice

Range in a Hyperbolic Lattice15. The range of any hyperbolic navigation system will largely depend upon the frequency andthe power of the transmitted signal. Generally, the lower the frequency the greater the range.Remember that, in order to double the range, it is necessary to quadruple the output power of thetransmitter. Hyperbolic navigation systems are inherently medium or long range systems (VOR/DME is a much more user friendly and light weight option for line of sight radio navigation) andtherefore operate with high power transmitters in the low frequency band.

16. The inability of the aircraft receiver to distinguish between ground waves and skywaves maylimit the useful range of a hyperbolic system.





17. Another limiting factor concerning useful range is system geometry. This is a problem whichoccurs when two hyperbolae which are being used to fix the aircraft position cross at an acute angle,thereby degrading the accuracy of the fix to an unacceptable level.





Loran C


Additional Secondary Phase Factors (ASF)

Transmission Characteristics

Indexing (Cycle Matching)

Presentation of Information

Skywaves

Chain Geometry

Range

Accuracy

Sources of Error



Loran C


15Loran C

1. Loran is an acronym for Long Range Aid to Navigation. Loran C is a hyperbolic navigationsystem but, unlike Decca, does not use the phase of signals to derive position lines. Instead, Loran Cuses a pulse transmission system and derives hyperbolic lines of position by measuring the timeinterval between pulses. The principle of operation is therefore said to be a differential range bypulse technique.

2. Loran C operates in the LF band using a carrier wave frequency of 100 KHz. At thisfrequency surface attenuation is low and the surface wave diffracts at an optimum rate, giving usablesurface wave signals at ranges of up to 1000 nm.

3. The surface wave is the primary propagation path for Loran C. Skywaves can also be used,but now the accuracy of the system is degraded.

Principle of Operation4. Consider Figure 15-1 which shows the familiar arrangement of a master transmitter (M) anda slave transmitter (S) of a hyperbolic system with baseline and baseline extensions.



Loran C


FIGURE 15-1Loran C Hyperbolic Lattice Construction

5. The master transmits a pulse omnidirectionally and this pulse is received a short time later atthe slave station. It is the reception of this (master) pulse which triggers the slave to transmit a pulseof its own, also omnidirectionally, but only after a fixed time delay. Let us call this delay Dmicroseconds (D µ sec).

6. Because the master and slave are a known distance apart, the propagation time of the original(master) pulse from M to S is also known. Let us call this time T microseconds (T µ sec).



Loran C


7. Consider now an aircraft at the master transmitter (M). The pulse transmitted from M willbe received instantaneously at the aircraft. The pulse will however take T µ sec to travel to S, therewill be a further D µ sec before the slave transmits and the slave pulse will then take T µ sec to returnto the aircraft. Thus, at the aircraft which is at M, the time interval between the reception of themaster pulse and of the slave pulse will be (2T + D) µ sec. With a little thought it should be obviousthat an aircraft anywhere on the master baseline extension will also receive the M and S pulses(2T + D) µ sec apart.

8. Consider now an aircraft at the slave transmitter (S). A pulse from the master transmitter willbe received, both by the aircraft and the slave, after T µ sec. The master pulse will of course trigger apulse from the slave (after the fixed time delay D) and this will be received instantaneously at theaircraft. Thus, at the aircraft which is at S, the time interval between the reception of the masterpulse and of the slave pulse will be D µ sec. Again with a little thought it should be obvious that anaircraft anywhere on the slave baseline extension will also receive the M and S pulses D µ sec apart.

9. For an aircraft anywhere other than at the master or slave stations, or on the baselineextensions, the time interval between the M and S pulses will be somewhere between (2T + D) and Dµ sec. Appreciate that the fixed time delay (D) ensures that the master pulse will always be receivedat the aircraft ahead of the slave pulse, regardless of the position of the aircraft within the chain.Return to Figure 15-1. An aircraft at position P will receive the pulses at the same time interval asany other aircraft positioned on the hyperbola passing through P. The hyperbola is therefore said tobe a line of constant time difference, and on a Loran C chart would be given a value which is the timeinterval (measured in µ sec).

10. Clearly, in order to fix the aircraft's position, a second position line is required and thus asecond slave is necessary.



Loran C


11. Each chain of stations in Loran C consists of a minimum of a master and two slave stations.Where geography permits, up to four slaves may be used. Slaves are designated W, X, Y and Z.

12. The timing of the slave transmission may be triggered by the reception of a pulse from themaster station, as previously described, or alternatively may be controlled by an atomic clock toachieve the same end.

Additional Secondary Phase Factors (ASF)13. Loran C receivers measure time delay differences in the signals they receive, computedifferences of distance, and hence determine the user’s position. The conversion from time to distancerequires knowledge of the signals’ velocities which differ from seawater values when propagatingover land. In addition to the signal delays experienced when propagating over smooth terrian andacross coastlines, extra delays are experienced where the path crosses elevated ground, especiallymountains.

14. Precise positioning requires the delays of land paths to be accurately mapped and theresulting data, in the form of additional secondary factors (ASFs), to be stored in Loran receivers.The ASF corrections are applied by the Loran receiver when calculating time delay values.

Transmission Characteristics15. The master and slaves of each Loran C chain do not transmit single pulses, but rather pulsetrains consisting of 8 pulses each separated by an interval of 1000 µ sec. These pulse trains aretransmitted approximately ten times a second. The master pulse train is distinguished from the slavepulse trains by means of a ninth pulse, which is transmitted 2000 µ sec after the eighth pulse. Theninth pulse in the master pulse train is made to blink at set rates in the event that there is anyunserviceability within the ground stations comprising the chain in question.



Loran C


16. In a chain consisting of a master and more than one slave, the delays (D) which are allocatedat each slave are so arranged such that the pulse trains will always arrive at the aircraft in a pre-determined order, regardless of where the aircraft is located within the area of coverage offered bythat chain. Thus, since the master station pulse train is easily identified by the ninth pulse, the pulsetrains from the slave stations are identified by the order in which they arrive at the aircraft.

17. The reason that pulse trains, rather than individual pulses, are used in Loran C is so that thepulse trains can be summed by the airborne Loran receiver. Summing enables the receiver to operatewith much lower signal to noise ratios and therefore increases the effective range of the system.

18. Mutual interference between chains is avoided because each chain is allocated a specific pulsetrain repetition period (remember that we said previously that the pulse trains were transmittedapproximately ten times each second). With Loran, the pulse train repetition period is termed thegroup repetition interval (GRI) and each chain has a unique GRI.

Indexing (Cycle Matching)19. The overall accuracy of a Loran C system depends, among other things, on the airbornereceiver's ability to measure accurately the time interval between each master pulse and theassociated slave pulse. In order for the airborne receiver to establish accurately this time interval acommon reference point within each pulse must be used. This is achieved by means of a processtermed indexing or cycle matching.

20. Each pulse within a pulse train (whether it be a master or a slave pulse train) has a pulsewidth of between 180 and 270 µ sec and therefore contains at least 18 complete cycles of carrierwave as shown at Figure 15-2.



Loran C


FIGURE 15-2Indexing/Cycle Matching in Loran C

21. Note that the pulse does not in fact have a square shaped envelope (which would be ideal) butstarts with a small amplitude, builds to a peak at about the eighth cycle and then diminishes. Itwould be advantageous to use the eighth cycle as the reference point within the pulse, however this isnot possible since, under certain circumstances, the surface wave pulse may by the eighth cycle bedistorted by the initial cycles of the same pulse which has arrived via a skywave propagation path.As usual it is necessary to compromise, and it is in fact the start of the third cycle which is used as theindex point within the pulse.

22. Each cycle within the Loran pulses represents a time duration of 10 µ sec. Providing thecorrect cycles of the Master and slave signals are used for time difference measurements, the accuracyof those measurements should therefore be within ± 10 µ sec.

23. If you are to have any hope of understanding how a Loran C receiver works it is importantthat you do not confuse the cycles of electromagnetic energy which comprise each pulse with thepulses which themselves comprise the pulse trains.



Loran C


Presentation of Information24. The control and display units (CDUs) of early Loran C receivers included a cathode ray tubeon which the pulses (as shown at Figure 15-2) were displayed. The task of the operator was tooverlay the slave pulse on the master pulse such that the start of the third cycle of both pulses wereprecisely aligned. In fact the two pulses were generated on successive timebases, and therefore thetime interval between the generation of the first and the second timebase was precisely equal to thetime interval between the index points of the pulses, providing that these were accurately aligned.This time interval was read from a LED display and the process then repeated for at least one othermaster/slave pairing within the same chain. The two (ideally three) time intervals were then plottedon the Loran chart (remember that the hyperbolae on a Loran chart are lines joining points of equaltime interval and that the hyperbolae are labelled with this time interval). The latitude and longitudeof the fix was then noted and the fix position transferred to a more suitable plotting chart.

25. With modern receivers the cycle matching and time interval determination is doneautomatically and therefore the cathode ray tube is no longer evident. The basic display of positionis as a latitude and longitude, however the receiver is normally interfaced with a navigation computercontaining pre-stored routes and waypoints. Now the navigational data available to the operatorwill include distance and time to the next waypoint, and so on. The outputs from such a sytem willbe very similar to those available from an Inertial Navigation System.

26. The modern Loran C receiver is designed to operate at low signal to noise ratios. In the eventthat an irregularity or malfunction occurs within a chain, the stations within that chain will transmitwarning signals which will typically be displayed as ‘blinking’ displays on the CDU.



Loran C


Skywaves27. Because the power output of Loran C transmitters is high, skywaves are present by day aswell as by night. Although the first part of the first pulse is sky wave free (because the sky wave takeslonger to arrive) the remaining pulses may suffer from sky wave interference. Skywaves are removedto some extent by varying the phase of individual pulses so that additional (skywave) pulses whichare received out of sequence will be out of phase and can therefore be eliminated.

28. At extreme ranges the surface wave pulse is very weak and eventually becomes unusable. Thesystem may still be able to produce an output of position by establishing time differences betweenskywave pulses rather than surface wave pulses. The resulting output is much less accurate and anautomatic system would warn the operator accordingly.

Chain Geometry29. Transmitter siting is largely dependent on geography. Maximum baseline length is about1000 nm. Coverage is widely available with chains located along the coastal areas of the NorthAtlantic, in the North Pacific and on the Hawaiian Islands, in the Mediterranean and in the MiddleEast.

Range30. Using surface wave propagation the range of this system is about 1000 nm. Skywavematching can be achieved to about 2000 nm. These ranges are from the master transmitter which isnormally at the centre of a chain.



Loran C


Accuracy31. As in any hyperbolic system, accuracy depends on the geometry of the chain. Quotedaccuracy is 0.25 nm or better (on 95% of occasions) in areas of good cover (using surface waves)reducing to about 1 nm at 1000 nm range from the master. At longer ranges, using skywaves,accuracy is poor but ± 10 nm at 1500 - 2000 nm range may be achieved.

Sources of Error32. The following errors are considered to be the major causes of inaccuracy when modern LoranC receivers are used to establish an aircraft's position. With the older airborne receivers operatorerror was a common, both because of misidentification of the third cycle during indexing and alsobecause of inaccuracies in the transposition of time difference on to the chart.

Static Interference33. Static will degrade the system since it transmits at 100 KHz. Remember that the lower thefrequency the more that static becomes a problem.

System Geometry34. There are few areas within the coverage of individual chains where the hyperbolae cross atoptimum angles, that is to say at 90° when two position lines are considered and at 60° when threeposition lines are used.






Basic Principles

Space Segment

User Segment

Control Segment

Accuracy

Sources of Error

Receiver Capability

‘All in View’ Receivers

Integrity Monitoring

Enhancement of GPS Information

Integrated Navigation Systems Using GPS

Future Applications of GPS





Applications of GPS - Area Navigation (RNAV) Approach Procedures





16Satellite Navigation Systems

1. Two Satellite Assisted Navigation (SATNAV) systems are currently operating. The NavstarGlobal Positioning System (GPS) is a constellation of 24 satellites operated by the US Department ofDefence (DoD) but available for civil use. The former Soviet Union (now Commonwealth ofIndependent States - CIS) also operates a similar system, called GLONASS (the global orbitingnavigation satellite system). In common with most other developments in modern technologyacronyms are used freely and these systems may also be referred to as GNSS (global navigationsatellite systems).

2. A brief comparison of the main parameters of the GPS and GLONASS navigation systems isshown in Figure 16-1:





FIGURE 16-1GNSS Parameters

3. One of the most important differences between GPS and GLONASS is that the latter does nothave any facility for degrading accuracy for unauthorised users. (see later)

Basic Principles4. The basic principle used in GNSS is that of determining range from the timed travel of radioenergy. With very accurate timing the range of the receiver (the aircraft) from the transmitter (thesatellite) can be established with minimal error.

5. In order to fix the aircraft position, ranges from two satellites must be established. Rangefrom a third satellite will confirm the accuracy of the fix and from a fourth satellite will enable bothvertical and horizontal position to be established (see Figure 16-2).

GPS GLONASS

No of Satellites 24 24

No of Orbits 6 3

Inclination of Orbits to Equator 55° 60°

Duration of Orbit approx. 12 hours approx. 11¼ hours

Satellite Altitude 20200 km 18840 to 19940 km

Signals L1 1575.42 MHzL2 1227.6 MHz

L1 approx. 1600 MHzL2 approx. 1250 MHz

Codes L1 P+C/A codesL2 P code

L1 P+C/A codesL2 P code





6. The aircraft's receiver assesses signal strength together with the relative positions of thesatellites when electing which of the available satellites (those that are above the horizon) to use inthe navigation solution. Precise timing is obtained by using an atomic clock to control the satellite'stransmissions whilst an electronic clock in the aircraft receiver controls receiver timing. Part of thesatellite's transmission includes details of its own position on its orbital path around the Earth, andthe time of the transmission.

7. The four basic information elements provided by GPS are:

(a) Latitude

(b) Longitude

(c) Height

(d) Time





FIGURE 16-2GNSS Principle of Operation





8. For the remainder of this chapter it is the GPS navigation system that is discussed, exceptwhere stated otherwise. There are 3 GPS segments :

(a) Space segment

(b) Control segment

(c) User segment

Space Segment9. The constellation of satellites in GPS consists of 24 operational satellites. The satellites areplaced in six circular orbital planes (3 or 4 per plane) all of which are inclined to the equator at 55°,as illustrated at Figure 16-3. The satellites orbit at an altitude of approximately 10,898 nm (20,200km) taking 12 hours to complete each orbit. The constellation is arranged so that at least 4 satellitesare within (line of sight) range of a receiver anywhere on Earth at any time. GLONASS operates in asimilar way with slightly different parameters.





FIGURE 16-3GPS Satellite Constellation

10. Two levels of navigation accuracy are provided by GPS : the Precise Positioning Service (PPS)and the Standard Positioning Service (SPS). PPS is a highly accurate positioning, velocity, and timingfacility which is made available only to authorised users, whereas SPS is a less accurate serviceavailable to all GPS users.





11. The main GPS satellite transmission frequencies are 1575.42 MHz (known as L1) and 1227.6MHz (known as L2). Under the old radar frequency bands nomenclature these are L bandfrequencies within the UHF band. Both frequencies are derived from the satellite's atomic clockfrequency of 10.23 MHz.

12. The L1 frequency is available for civil use for the provision of navigation fixing. The L2frequency is used exclusively for military navigation and can be only be accessed by means of specialuser codes. The benefit of the L2, ’P’ code signal (described later) is that it allows users to correct forionospheric propagation delays. Some receivers are able to measure the delay between the signal inthe L1 frequency and the L2 frequency without access to the P code. There are plans to add, in futuresatellites, another frequency for civil users so they can easily correct for ionospheric delays.

Precise Positioning Service (PPS)13. The PPS is primarily intended for military users and the authorisation for its use is decided bythe US DoD. A selective availability (SA) feature can be activated to reduce the GPS accuracy tounauthorised users by introducing controlled errors into the signals. The level of degradation can bevaried and is accomplished by intentionally ‘dithering’ the satellite clocks and ephemeris information(see later).





14. An anti-spoofing (A-S) feature is also introduced at random times without warning toinvalidate any copying of PPS signals by unauthorised individuals. The technique alters the P codecrytographically into a code denoted as Y code. Encryption keys and techniques are provided to PPSusers which allow them to remove the effects of the SA and A-S features. PPS receivers can use eitherthe P(Y) or the C/A code (see later) or both. Maximum accuracy is obtained using the P(Y) code onboth L1 and L2. The difference in propagation delay between the two frequencies is used to calculateionospheric corrections. P(Y) capable receivers commonly use the C/A code initially to acquire GPSsatellites and determine the approximate P(Y) code phase although some P(Y) receivers are able toacquire the P(Y) code directly by using the precise clock.

Standard Positioning Service (SPS)15. The SPS is primarily intended for civilian use and is specified to provide a 30 metre horizontal positioning accuracy to any GPS user in peacetime. The accuracy specification would bereduced to ± 100 metres in the event that the selective availability feature is activated. Typically aSPS receiver uses only the C/A code and an ionospheric model to calculate corrections, which is a lessaccurate technique than measuring dual frequency propagation delays. The accuracy specificationalso includes any ionospheric modelling error.

(2σ)





Transmissions – Characteristics and Timing16. The principle of obtaining range position lines is relatively simple and is of course theprinciple upon which DME operates. However, with DME the position of the source of the range arcis known, and the start of the transmission timing is known because it originates at the aircraft. In aSATNAV system the aircraft equipment does not transmit and so some other method is necessary toestablish the time of travel of the satellite's transmission. In addition, with SATNAV, the origin of theinformation is continually moving, since the satellites are all orbiting the Earth . Digital processingtechniques are clearly necessary to co-ordinate and analyse the large quantities of data which isneeded to resolve these problems.

17. Time is computed in Space Vehicle (SV) time, GPS time and UTC. SV time is the timemaintained by each satellite. Each SV contains four atomic clocks (2 cesium and 2 rubidium) which

control the satellites’ transmissions to a timing accuracy of better than (1 nano-

second). SV clocks are monitored by the master control ground station (MCS) and are occasionallyreset to maintain time to within 1 millisec of GPS time. SV clock correction information is passedfrom the MCS, via satellite to the user, to correct for any residual SV timing error.

18. Since the aircraft equipment also needs to know the start time of the transmission it requiresits own clock. An electronic clock is provided in the aircraft receiver/processor and this is lessaccurate than the atomic clock which is used in the satellite, however this problem is overcome in thefixing process. It would not be feasible to incorporate an atomic clock into the aircraft equipment ona cost basis alone. SV time is passed to the GPS receiver, as part of the GPS signals, and it is thenconverted to GPS time.

11 000 000 000, , ,--------------------------------------- sec





19. GPS Time is measured in weeks and seconds from 24:00:00, January 5, 1980 and is steered towithin one microsecond of UTC. GPS Time has no leap seconds and is ahead of UTC by severalseconds (currently approximately 13 seconds). Time in Universal Coordinated Time (UTC) iscomputed from GPS Time using the UTC correction parameters sent as part of the navigationmessage.

20. Each GPS satellite generates a unique coded digital transmission. The codes transmitted onL1 are known as the C/A (coarse/acquisition) code, and the P (precision) code. The C/A code isrepeated continuously, a sequence being completed every millisecond. This period is called the epochand is used in civilian GPS receivers. The pattern of the digital bits that make up the C/A code is solarge (1023 bits) that it appears to be random. Hence it is referred to as a pseudo-random code andrequires a bandwidth (spectrum spread) of 1 MHz.

21. The P code element is even longer and the part of the code sequence allocated to each satelliteis re-initiated every 7 days. This code provides more precise navigation accuracy but access to it iscontrolled by the US DoD and, as with the navigation facilities provided by the L2 frequency, is notgenerally available for civil use.

22. In addition to the fixing transmissions each satellite also transmits a data bit stream at 50 bitsper second. This navigation message provides information on the satellite's orbital position andother status and correction factors. Each message takes 30 seconds to complete.

23. The precise format of the navigation message is illustrated at Figure 16-4.





FIGURE 16-4Navigation Message Format

User Segment24. The GPS user segment can perform two basic measurements of the GPS signals. The firstmethod compares the C/A or P code that it is receiving with a locally generated copy in order tocompute the transmission delay between the satellite and the receiver. This measurement is convertedto range and by using four or more satellites it is possible to determine the position of the user oncethe position of the GPS satellites has been obtained using the ephemerides of the navigation message.





25. The GPS receiver produces an internal digital code which is identical to the C/A codetransmitted by each satellite. The timing of the satellite transmission and the GPS receiver codeswould in an ideal case be perfectly synchronised. Thus the GPS receiver would recognise which partof the pseudo-random code it had just received and would then compare it with the same part of theinternally produced reference code. The time difference between the two points, at a speed of 3 x

108 m/sec, gives the distance that the signal has travelled. This distance is the range of the aircraftfrom the satellite. Figure 16-5 illustrates this process diagrammatically.

FIGURE 16-5Measurement of range using a binary sequence

26. The GPS receiver in the aircraft has a clock which is not as precise as the satellite clock andtherefore its interval code is not likely to be perfectly synchronised with the satellite C/A code. As aresult of this de-synchronisation, which is known as clock bias, the range calculation will be in error,and is therefore referred to as a pseudo-range (PR).

27. The pseudo-range is corrected to give true range by applying a correction for the clock bias.The range error due to timing inaccuracy alone is 0.3 metres for each nano-second error.





28. The second and more precise method of ‘ranging’ is to obtain the difference in phase anglebetween the received carrier signal and a receiver-generated signal at the same frequency. Thismeasurement is known as the carrier phase observable and it can reach millimetre precision.However, it lacks the accuracy of the pseudo-range because once the tracking is started, the phasecan only be identified with an ambiguity of an unknown number of times the carrier wavelength(about 19 cm for L1).

Signal Acquisition29. The satellite signal level near the Earth is less than the background noise (hence the GSPaerials should be positioned as near to the receiver as possible to avoid long cable runs) andcorrelation techniques are used by the GPS receiver to obtain the navigation signals. A typicalsatellite tracking sequence begins with the receiver determining which satellites are visible for it totrack. Satellite visibility is based on the user-entered estimates of present position, time and date andstored satellite almanac information. If no stored data exists or if only very poor estimates ofposition and time are available the receiver must search the sky in an attempt to locate randomly andlock on to any satellite in view; under such circumstances it may take up to 15 minutes for the GPSreceiver to obtain the first position fix. If the receiver can estimate satellite availability it will target asatellite to track. Once one satellite has been acquired and tracked the receiver can decode thenavigation message and read almanac information about all of the other satellites in theconstellation.

30. When a new GPS receiver is first switched on, it must download the almanac and ephemerisdata before it can determine position. This usually takes about 12.5 minutes.





Signal Tracking31. The receiver has a carrier tracking loop which is used to track the carrier frequency and acode tracking loop which tracks the C/A and P code signals. The two tracking loops work together inorder to acquire and track the satellite signals. The receiver’s carrier tracking loop generates a carrierfrequency which differs from the received carrier signal due to Doppler shift. This Doppler shift isproportional to the relative velocity along the line of sight between the satellite and the receiver.

Position Fixing32. Fixing is achieved using the range principle as illustrated at Figure 16-6. Having firstestablished the satellites’ position and velocity the GPS receiver starts to acquire a navigationsolution, after which the GPS enters the navigation mode. In this mode several pseudo rangemeasurements (typically seven) are taken from all the satellites within view, and this process isrepeated once every second. Since the airborne GPS receiver’s clock will usually be in error, thepseudo range measurements will all be too short or too long by the same amount. With pseudoranges from four satellites, there are four simultaneous equations to solve the common error, thusremoving clock bias and producing a very accurate three-dimensional fix. Figure 16-7 shows a three-satellite fix for clarity.





FIGURE 16-6Position Fixing in GPS





FIGURE 16-7Correction of User clock bias in GPS

33. Initially, conversion of pseudo-range measurements is carried out to find the users position interms of an Earth Centred/Earth Fixed (ECEF) co-ordinate system. In the ECEF system any positioncan be defined in terms of 3 axes (X,Y,Z), which originate at the centre of the spherical Earth, asshown at Figure 16-8.





FIGURE 16-8Earth Centred/Earth Fixed Co-ordinate System (ECEF)





34. In Figure 16-9, 4 satellites are being monitored by a user. The satellite positions (X1Y1Z1,

X2Y2Z2, X3Y3Z3, X4Y4Z4) are given in the satellite message and it is, therefore, a relatively simple

calculation, using the 4 pseudo-ranges, to determine the position of the user receiver (X5Y5Z5) in thesame ECEF co-ordinates. The alogorithm used in this calculation has, in fact, four unknowns andtherefore not only is it possible to determine the users X,Y,Z co-ordinates but clock bias can also becalculated. (See Paragraph 32).





FIGURE 16-9Determination of User Position in ECEF Co-ordinates





35. Using transformation equations, the user receiver ECEF co-ordinates are converted togeodetic co-ordinates of position and altitude over a reference ellipsoid (see Figure 16-10).An input of pressure altitude, if applicable, is also used to compare continuously with the GPSvertical position calculations.

FIGURE 16-10Conversion of ECEF Co-ordinates to Geodetic Co-ordinates

φ λ h, ,( )





GPS Geodetic Datum36. Throughout the evolution of GPS there has been a need for a rationalisation of the datumused as the basis of navigation using the system. That is to say that local datums exist which usedifferent ellipsoids that approximate the shape of the Earth over a selected area, but which are notvalid on a global scale. Conversion between datums is possible, however inherent inaccuraciespresent in National (local) datums can result in large residual errors.

37. The absolute datum which has now been adopted by ICAO, the United kingdom andEurocontrol is WGS-84 which stands for the World Geodetic System 1984. This is the singlegeodetic reference system for civil aviation effective 1 January 1998.

Control Segment38. The GPS control segment tracks and monitors the signal from the GPS space segment.Stations located around the world monitor the performance of the satellites (e.g. orbits and clockbehaviour), and a master control station in Colorado Springs sends up corrections for transmissionto the users, at least once per day, if errors are detected.

Accuracy39. The position accuracy of the C/A code signal is potentially better than ± 30 metres.

Sources of Error40. The resolution of clock error (or clock bias) has already been described. Other sources oferror are now considered.





Satellite Geometry Error41. The general term used to describe this error in GPS is GDOP (geometric dilution of precision).This error occurs when two satellites are close together and the angle of cut between range circles isshallow. Having orbital planes at 55° (GPS) and 60° (GLONASS) to the equator helps to reduce thisproblem. Errors of between 30 and 70 metres can nevertheless result when the problem occurs. Insome GPS receivers it is possible to select alternative satellites in order to improve GDOP, however itshould be noted that fixing continuity will be interrupted for a short period during the changeover.GDOP can be further sub-divided as shown in Figure 16-11.

FIGURE 16-11Types of Dilution of Precision in GPS





Satellite Ephemeris Error42. The orbit parameters of the satellite are referred to as ephemeris. Errors in satellite positiondirectly induce fixing errors. The master control ground station (MCS) monitors and controls thesatellite's orbital position to ensure that range errors resulting from ephemeris inaccuracies remainwithin ±.5 metres.

Satellite Clock Error43. Transmission timing and range measurement is dependent on the satellite clock. Clock errorsare corrected by the MCS. This range error resulting from clock inaccuracies should not exceed ± .5metres.

Ionospheric Error44. This error is also known as atmospheric propagation error. The signals passing downthrough the ionosphere are slowed by the small amount of refraction that occurs. The error will varywith time of day, year and elevation of the satellite but should not be more than 4 metres.

Multipath Error45. Multipath is caused by reflected signals from surfaces near the receiver that can eitherinterfere with, or be mistaken for the signal that follows the straight line path from the satellite.Multipath is difficult to detect and sometimes hard to avoid, however, the effects can be mitigated toa certain extent by special antenna design and/or enhanced receiver software. Any range errorresulting from multipath propogation should not exceed ± .5 metres.





Orbital Perturbations46. It was stated that GPS satellites are positioned in circular orbits around the Earth;unfortunately, in reality, each satellite is subject to several major influences which distort that orbit.The major problem is caused by the earth’s equatorial bulge, however solar wind, the gravitationalpull of sun/moon/planets, and other influences all effect the GPS system. Corrections for GPS orbitalperturbations are defined and broadcast at least once per day, via the satellites to the user, along withclock correction data.

Instrument/Receiver Error47. Errors can arise in the GPS receiver due to electrical noise as well as in time measurement andrange/position computation. Range errors of 1 metre are possible due to such causes.

Receiver Capability48. The speed of operation and the accuracy of the aircraft SATNAV receiver will naturallydepend on the complexity, and therefore the cost and the weight, of the equipment which is installed.There are basically three types of receiver which may be used.

Sequential Tracking Receivers49. A sequential receiver tracks the necessary satellites by using one or two channels. The set willtrack one satellite at a time and combine all four pseudorange measurements once they have beenmade. These receivers are amongst the cheapest available but cannot operate under high dynamicsituations and have the slowest time to first fix.





Multi -Channel Receivers50. A continuous tracking receiver must have at least four channels in order to track foursatellites simultaneously. GPS receivers are available with up to 12 channels, but due to their greatercomplexity multiple channel sets involve proportionally higher costs. Four and five channel sets offersuitable performance and versatility, tracking 4 satellites simultaneously; a five-channel receiver usesthe fifth channel to read the navigation message of the next satellite to be used when the receiverchanges the satellite selections. The fifth channel is also used in conjunction with each of the otherfour for dual frequency measurements. A multi channel receiver is the best for high dynamic vehiclessuch as aircraft.

Multiplex Receivers51. A multiplex receiver switches at a fast rate (typically 50 Hz) between the satellites beingtracked, continuously collecting sampled data to maintain two to eight signal processing alogorithmsin software. The navigation message data is read continuously from all the satellites.

‘All in View’ Receivers52. GPS receivers traditionally choose the four satellites of those available which give the bestgeometry to perform a position fix. However, in situations where one or more of the satellites aretemporarily obscured from the antenna’s view, the receiver will have to acquire additional satellitesignals to generate a solution. The accuracy will therefore degraded until the new satellite is acquired.In order to overcome this problem a receiver can be designed to use all available satellites in view,typically six or seven, to generate the solution. If one or two satellites are then lost from view therewill be little or no loss of accuracy. The receiver will need a channel for each satellite or will have touse multiplex techniques.





Integrity Monitoring

Receiver Autonomous Integrity Monitoring (RAIM)53. This function enables the aircraft GPS receiver to monitor the integrity of incoming satellitesignals, to identify the satellite responsible for an erroneous signal, and to remove that satellite fromthe navigation solution.

54. We know that the aircraft receiver uses the signals from four satellites with a suitablegeometry to establish the aircraft's horizontal and vertical position. There will however be morethan four satellites above the horizon, and it is the signals from these ‘spare’ satellites which are usedto achieve the RAIM function.

55. The situation is as follows:

(a) A fifth satellite signal enables the receiver to establish that one of the four signalsbeing used to determine the aircraft's position (horizontally and vertically) iserroneous.

(b) A sixth satellite signal enables the receiver to identify the satellite which is giving theerroneous signal and to remove it from the navigation solution. This is called FaultDetection and Exclusion (FDE).

(c) In some airborne systems the GPS receiver uses barometric altitude as anaugmentation to RAIM. In other words the output from the pressure altimeter or themode C of the SSR is compared with the altitude determined by the GPS receiver. Inthis case, providing that the geometry of the satellites is satisfactory, the number ofsatellites required to perform the full RAIM function is reduced from six to five.





(d) It is worth noting that Integrity messages need not necessarily be generated by a userreceiver, they may also be transmitted from an earth station or communicationsatellites.

56. In airborne equipment where a GPS receiver provides data to an integrated navigation system(a flight management system or a multi-sensor navigation system), either the GPS receiver must beproviding RAIM or the multi-sensor navigation system must be providing a level of integrity which isequivalent to that given by RAIM, before the GPS can be used as the primary navigation referencefor flight under IFR.

57. In airborne equipment comprising a stand alone GPS receiver, the RAIM must be functioningbefore the GPS can be used as the primary navigation reference for flight under IFR.

58. Not all GPS receivers are RAIM equipped. For those receivers that are RAIM equipped, it isnot always the case that six satellites are available with an appropriate geometry and each with asufficient elevation above the horizon. The RAIM function may therefore be interrupted.

59. The limitations discussed above make GPS suitable for use only as a supplemental airnavigation system for certain phases of flight. The use of GPS in any form for any type or any part ofa precision approach is not permitted by the UK CAA.

Aircraft Autonomous Integrity Monitoring (AAIM)60. AAIM is an ’integrated’ system where individual navigation systems cross check each otherthrough the navigation computer. The Inertial Navigation System derived position can be comparedto the GPS position and to positions derived from other navaids. A barometric or radio altitude canbe used to cross-check the GPS height output in a three-dimensional fix.







Within a particular hyperbolic lattice between a Master and a Slave station, where does the area ofambiguity lie:

QUESTION 2.

Loran C operates at a frequency of:

QUESTION 3.

Additional Secondary Phase Factors are Loran C corrections which compensate for:

QUESTION 4.

The signal from a Loran Master station is identifiable from those from the slave stations because:

QUESTION 5.

Cycle matching/indexing in Loran is:

QUESTION 6.

Loran position fixing accuracy is quoted as:





QUESTION 7.

The biggest advantage of the GLONASS satellite system over GPS, for commercial aviation, is that:

QUESTION 8.

The GLONASS satellites have an orbital altitude of:

QUESTION 9.

What are the 4 basic information elements provided by GPS:

QUESTION 10.

As a civilian user of the GPS system it is possible to make use of the _______________ service.

QUESTION 11.

How is SV time produced in the GPS system:

QUESTION 12.

The GPS Navigation message comprises a total of ______ subframes.

QUESTION 13.

When a user is near the earth, how does the received satellite signal level in GPS compare in relationto the background noise.

QUESTION 14.

Why do the GPS P code signals change at random intervals to become Y code signals:





QUESTION 15.

Is GPS time the same as UTC:

QUESTION 16.

How long does it take for a brand new GPS receiver to download almanac and ephemeris data froma satellite following initial switch-on:

QUESTION 17.

How many satellite signals does a GPS receiver need to receive in order to correct for User ClockBias:

QUESTION 18.

Which Geodetic Reference System has been adopted to convert GPS position information intolatitude and longitude:

QUESTION 19.

What are the three GPS segments:

QUESTION 20.

Which of the two GPS signals, L1 or L2, provides both P code and C/A code information:





ANSWERS:ANSWER 1.

The area of ambiguity is in the area immediately surrounding the baseline extension

ANSWER 2.

Loran C operates at 100 Khz

ANSWER 3.

Additional Secondary Phase Factors are Loran corrections that compensate for variations in radiowave velocity, due to the type of surface over which the wave is travelling

ANSWER 4.

The Master station transmits a total of nine pulses (the slave stations transmit eight pulses); the ninthpulse is double-spaced from the previous eight

ANSWER 5.

Cycle matching is the process by which the equipment identifies the third cycle of each pulse, beforemeasuring the time difference between the two

ANSWER 6.

Groundwave ±1 nm at 1000 nm range, skywave ±10 nm at 1500-2000 nm range.

ANSWER 7.

The GLONASS system does not have an accuracy degredation facility





ANSWER 8.

The GLONASS satellites orbit the earth at an altitude of approximately 19000 km

ANSWER 9.

The 4 basic information elements provided by GPS are latitude, longitude, height and time

ANSWER 10.

A civilian user can only make use of the Standard Positioning Service (SPS)

ANSWER 11.

Each satellite has 4 atomic clocks which control satellite transmissions to an accuracy better than 1nano-second

ANSWER 12.

The GPS Navigation message comprises 5 subframes

ANSWER 13.

The signal level is less than the background noise and requires correlation techniques in order toobtain the navigation data

ANSWER 14.

The Y code is produced at random, using cryptographic techniques, to provide an anti-spoofingfacility which invalidates any copying of the PPS signal by unauthorised persons.





ANSWER 15.

No, GPS time is measured in weeks and seconds from 24:00:00 on 5 January 1980

ANSWER 16.

It takes 12.5 minutes to download almanac and ephemeris data from a satellite to a new GPS receiver

ANSWER 17.

A GPS receiver needs 4 satellite signals in order to correct for User Clock Bias

ANSWER 18.

GPS receivers use the WGS 84 Geodetic Reference System to calculate latitude and longitude

ANSWER 19.

The three GPS segments are the Space segment, the Control segment and the User segment

ANSWER 20.

The L1 signal contains both P code and C/A code information

Enhancement of GPS Information61. Apart from errors due to GDOP the only significant error in GPS is Ionospheric error.Consideration has been given to using GPS for precision approach purposes but with the basicsystem accuracy this concept is untenable. However, ways around the problem are being researchedfor application on a more localised basis, in order to achieve accuracy figures within 1 to 3 metres.This sort of figure compares favourably with the ILS Cat 1 vertical accuracy requirement of 4 metres





at the threshold. Aspects such as the time taken to detect and alert the pilot to system accuracydegradation are also being addressed.

Differential GPS62. The first method of GPS enhancement is what is termed Differential GPS. The principle ofthis technique is that a ground station's GPS derived position can be compared to its exact, knownposition. The difference between the two must be due to the GPS fixing error (principallyionospheric error). A correction (data correction message), termed the differential correction to GPSposition is passed via a data link such as ACARS (aircraft communications addressing and reportingsystem) to all suitably equipped aircraft which are within range. The aircraft's GPS position outputis thus corrected to give the highest accuracy possible. Figure 16-12 is a simple illustration of themechanics of Differential GPS.





FIGURE 16-12Differential GPS - Principle of Operation





63. There are two main types of Differential GPS in use : Local Area Differential GPS (LADGPS)and Wide Area Differential GPS (WADGPS). LADGPS works in the manner described in Paragraph62 with a reference station in the vicinity of say, an aerodrome, providing communication direct tothe relevant aircraft. WADGPS uses ground reference stations spaced hundreds of miles apart feedinga master control station which sends correction data via geostationary communication satellites ;with this latter system there are a number of ground reference stations possibly stretching across anentire continent.

64. Note that aircraft using corrected satellite information from a Differential Ground Stationwill have satellite integrity information automatically passed to them if necessary.

Wide Area Augmentation System (WAAS)65. Basic GPS fails to meet the accuracy, integrity and availability required by an airborne user.The WAAS was developed as a way of improving all 3 aspects of the basic system.

66. The WAAS network that has been developed in the USA comprises approximately 35 groundreference stations that cover a very large service area (the entire USA plus parts of Canada andMexico).

67. WAAS consists of the following components :

• Wide-Area Reference Stations

• Wide-Area Master Stations

• Ground Earth Stations

• Geostationary Satellites





68. Signals from GPS satellites are received by the wide area ground reference stations (WRSs).Each of these precisely surveyed reference stations receive GPS signals and determine if any errorsexist. These WRSs are linked to form the WAAS network. Each WRS in the network relays the datato the wide area master station (WMS) where correction information is computed. The WMScalculates correction alogorithms and assesses the integrity of the system. A correction message isprepared and uplinked to a Geostationary Communication Satellite via a ground uplink system(GUS). The message is then broadcast on the same frequency as GPS (L1, 1575.42MHz) to receiverson board aircraft which are flying within the broadcast coverage area of the WAAS. Thecommunications satellites also act as additional navigation satellites for the aircraft, thus, providingadditional navigation signals for position determination.

69. The WAAS will ;

(a) improve basic GPS accuracy to approximately 7 meters vertically and horizontally.

(b) improve system availability through use of the geostationary communication satellitescarrying the necessary equipment to generate ‘GPS like’ navigation signals.

(c) provide important integrity information about the entire GPS constellation.Figure 16-13 illustrates the main features of a WAAS.

70. The USA WAAS network was expected to be released for initial operations during 1999however, approval for full operational capability is not now expected before 2002. A fullyoperational WAAS network will provide users with accurate position information sufficient forCategory 1 precision approaches.





FIGURE 16-13Wide Area Augmentation System





Local Area Augmentation System (LAAS)71. A second type of augmentation to the GPS signal is the Local Area Augmentation Systemdesigned for CAT II and III precision approaches. The LAAS is intended to complement WAAS andhence supply users with seamless satellite based navigation for all phases of flight. In practical terms,this means that locations where the WAAS is unable to meet existing navigation and landingrequirements (such as availability), the LAAS will be used to fulfill those requirements. In addition,the LAAS will meet the more stringent Category II/III requirements that exist at selected aerodromes.LAAS will also provide the user with a navigation signal that can give an all weather groundnavigation capability enabling the potential use of LAAS as part of a surface navigation system andinput to surface surveillance / traffic management systems.

72. The LAAS consists of the following :

(a) Local-Area Reference Stations

(b) Local-Area Central Processor Station

(c) Local-Area Data Transmitter

(d) Local-Area Pseudolites (if needed - see later)

73. Similar to the WAAS concept (which incorporates the use of communications satellites tobroadcast a correction message), the LAAS will broadcast its correction message via very highfrequency (VHF) radio datalink from a ground-based transmitter.





74. The LAAS ground reference stations will be installed at precise locations at an airport. Theground reference stations receive and collect GPS positioning data. The GPS data are then sent to thecentral processing station, which compares the calculated positions with the known precise locationsto determine the errors of the GPS. The GPS corrections are then transmitted to aircraft in thevicinity of the airport.

75. LAAS will provide the extremely high accuracy, availability, and integrity necessary forCategory II/III precision approaches. It is fully expected that the full configuration will pinpoint theaircraft’s position to within one meter or less. Figure 16-14 shows the key elements of a LAAS.





FIGURE 16-14Local Area Augmentation System





Pseudolites76. The various GPS enhancements that have been mentioned so far all rely on Satellite basedtransmitters for the production of GPS signals. In contrast it is also possible to use ground basedtransmitters that broadcast GPS-like signals to supplement those generated by the satellites. Theseadditional, ground-based transmitters were originally called ’pseudo-satellites’ which becameshortened to ’pseudolites’.

77. Compact in size, the ground-based low-power transmitters each fit entirely on a circuit boardthe size of a credit card. Capable of running on a 9-volt battery for over 12 hours, the inexpensivedevices transmit just a few microwatts of power, emulating a GPS satellite. The beacons are typicallysituated in pairs on either side of the approach path to the runway. Power of the broadcast signalsfrom the pseudolites is set low, measurable only inside a ‘bubble’ emanating from the transmitter.

78. The additional ranging signal from a pseudolite can be extremely useful. Each additionalpseudolite signal allows the user to perform basic navigation, fault detection, and fault isolationusing one less satellite signal than would otherwise be required. Also, significant improvements invertical position accuracy becomes possible, especially for applications in which pseudolites may beplaced below a vehicle to improve geometry (for example, for an aircraft on an approach).

79. To allow a stand-alone GPS receiver to use pseudolites as an extra code-based ranging source,a pseudolite signal’s timing accuracy must be comparable with that of the satellite signals.Incorporating a pseudolite with an inexpensive clock into a code-based differential GPS (DGPS)system provides a practical solution to the timing problem. The reference receiver synchronouslymeasures the pseudolite code-phase and transmits pseudorange correction information(incorporating the pseudolite clock error) to the pseudolite users. This procedure is basically identicalto satellite-based DGPS and fits readily into an existing differential system.





80. One of the first applications to use pseudolites was the Integrity Beacon Landing System(IBLS), shown in Figure 16-15. This positioning system comprises a ground reference system and avehicle navigation system. The IBLS ground system includes a carrier-phase DGPS (CDGPS)reference receiver, a datalink from the reference station to the aircraft, and a pair of integrity beaconpseudolites underneath the aircraft’s final approach path to the runway. The pseudolites transmitoverlapping, hemispherical signal ‘bubbles’ that the aircraft flies through just before landing. TheIBLS aircraft system includes a CDGPS user receiver and a computer to process the IBLSalogorithms. The user receiver has two antennas: one facing up to collect satellite signals and onefacing down to collect pseudolite signals.





FIGURE 16-15Integrity Beacon Landing System using Pseudolites





81. As the aircraft leaves the bubbles, its computer solves the alogorithms to check the integrity ofthe system. If the check exceeds a certain threshold, an error has occurred and the approach isaborted; this is the fundamental integrity monitor in IBLS. One can set this threshold so that theprobability of an undetected navigation failure satisfies almost any given integrity requirement. Theend result is a highly robust, maximum performance navigation system.

82. In a recent flight trial 110 autopilot landings of a Boeing 737 were completed using such asystem. The integrity beacons provided consistent accuracies in the order of a few centimeters duringeach of the automatic landings.

European Geostationary Navigation Overlay System (EGNOS)83. The European Space Agency, the European Commission and the European Organisation forthe Safety of Air Navigation (Eurocontrol) are jointly developing EGNOS, Europe’s augmentationsystem for satellite navigation. This ECU 150 million project will provide civil GPS or GLONASSusers with improved accuracy, integrity and availability.

84. Two transponders are being flown by two Inmarsat-III satellites, located at longitudes 64 East(Indian Ocean Region - IOR) and 15.5 West (Atlantic Ocean Region - East - AOR - E). Together theywill cover not only the whole of Europe but Africa, South America and most of Asia. The IORsatellite was launched on 3 April 1996, the AOR-E satellite was launched in August 1996.

85. In its final set-up, EGNOS will provide Ranging, Integrity and Wide Area DifferentialServices.

• The Ranging Service will broadcast GPS-like navigation signals to improve overall satellitenavigation service availability. For instantaneous determination of his position, a user has to





receive signals from four satellites. Neither GPS nor GLONASS can provide this at all timesand all locations worldwide. EGNOS will help to fill this gap.

• The Integrity Service will broadcast range error estimates for each GPS, GLONASS orEGNOS navigation signal. Without this EGNOS capability, information on abnormal perfor-mance or failure of GPS and GLONASS would take 15 minutes or longer to reach the user.The Integrity Service will enable users to decide whether a navigation satellite signal is out oftolerance before any critical situation arises.

• The Wide Area Differential Service will broadcast correction signals to improve the precisionof satellite navigation. The Wide Area Differential Service will establish a position fixingaccuracy of 5 - 10 metres.

86. The Ranging Service started in 1997. The other services are being introduced graduallybetween 1998 and 2003.

87. EGNOS itself will be composed of :

• The space segment : two INMARSAT III transponders, later to be extended to meet theextreme safety requirements for certain aircraft precision approaches to airports.

• The earth segment : Ranging and Integrity Monitors distributed over the service area will beconnected to Master Control Centres, where the EGNOS signals will be generated. At leastthree such centres are needed to meet civil aviation safety requirements. The France Telecom’searth station at Aussaguel and that of Deutsche Telekom at Raisting will be used as primaryaccess stations, respectively for the INMARSAT III, AOR-E and IOR navigation transpon-ders.

• The user segment : EGNOS standard receivers.





Integrated Navigation Systems Using GPS88. Integrated navigation systems are available which combine the outputs from a number ofindividual navigation equipments in order to provide an ultimate solution by statistically analysingall of the relevant data. Such systems are termed Multisensor Systems, two examples of which arediscussed below :

Combined GPS/GLONASS Receivers89. The development of combined GPS/GLONASS receivers provides an absolute positioningaccuracy which is comparable with that produced by the GPS L1 C/A code (without S/A mode). Thepotential accuracy figures for a combined GPS/GLONASS receiver are illustrated at Figure 16-16.





FIGURE 16-16Combined GPS/GLONASS Performance Characteristics

90. In summary, the use of GLONASS satellite signals in addition to GPS provides verysignificant advantages :

• twice the number of satellite signal observations

• reduction in the Geometrical Dilution of Precision factor

• no possibility of precision degrading in GLONASS (unlike GPS Selective Availability mode)

• unaffected by Anti-Spoofing measures which reduce the performance of GPS

GPS receiver,L1 frequency, C/A code (SPS),without S/A mode

GPS receiver, L1 frequency, C/A code (SPS), with S/A mode

GPS receiver, L1 frequency, C/A code (SPS) in differential mode

GLONASS receiver (absolute positioning)

Combined GPS/GLONASS receiver (absolute positioning)

Horizontal positioning accuracy

15-20 m 100 m 1-3 m 15-20 m 15-20 m

Vertical positioning accuracy

50 m 150 m 1-3 m 50 m 50 m

Velocity vector component accuracy

2-3 cm/s 1 m/s 2-3 cm/s 2-3 cm/s 2-3 cm/s





GPS and INS Integration91. Integrated navigation systems are available which can combine the outputs of an InertialReference System (IRS) with GPS. The position and velocity elements provided by the IRS arecontinually checked against GPS inputs and if GPS signals are unobtainable or unreliable, ‘modified’IRS data is used instead until such time as the GPS signals are regained. Such systems could, inaddition to providing the normal navigation facilities, also produce vertical navigation (V NAV)outputs providing at least 4 satellites are available.

92. Another way of looking at the benefits of the above system is as follows;

(a) The short term accuracy of an IRS is extremely good, however in the long term theposition error degrades with time.

(b) The long term accuracy/stability of GPS is good however there may be short periodswhere satellite signals are not available. By combining the two systems we benefitfrom both the short term accuracy of the IRS and the long term accuracy of GPS, thusgiving the optimum solution.

Future Applications of GPS

Enhanced Ground Proximity Warning System (EGPWS)93. The current generation of GPWS equipments suffer from two main problems :

(a) Short warning time to potential impact with terrain

(b) Minimal advice generated as to the best method of avoiding the hazard.





The development of EGPWS is an attempt to overcome these problems.

94. EGPWS operation depends upon a worldwide terrain database which is held in the EGPWScomputer memory. The system uses data provided by GPS in order to display surrounding terrainbelow, at or above the aircraft’s altitude. In brief, the system sounds an audible warning if theaircraft’s projected flight path takes it too close to terrain.

95. With a normal GPWS equipment pilots have no visual display to confirm terrain details,merely the alert and warning aural message/lights (typically 30 seconds warning) which identify thata hazard exists. The EGPWS system displays the surrounding terrain up to 320 miles away, from itsdatabase, and provides up to a 60 second warning of a potential impact.

96. The visual display can present terrain details on the aircraft’s weather radar display or EFISscreen in one of three colours, depending on proximity. Green terrain is below the aircraft, yellow isabove, and red is well above. Screen resolution also gets denser as the height of the terrain increases.If the system issues an alert, the terrain that poses a threat is shown as a solid block of yellow or red.

Automatic Dependent Surveillance Broadcast (ADS-B)97. ADS-B is technology that allows pilots in the cockpit and air traffic controllers on the groundto monitor aircraft traffic with much more precision than has previously been possible.





98. Each ADS-B equipped aircraft broadcasts its GPS position via a digital data link along withother data, including airspeed, altitude and whether the aircraft is turning, climbing or descending.This provides anyone with ADS-B equipment with a much more accurate depiction of air traffic thanradar can provide. Since the equipment is small and light, it can be made a standard part of theequipment onboard an aircraft, allowing pilots to see an accurate depiction of real-time air traffic inthe same manner as the controllers. ADS-B works at all altitudes and on the ground therefore it canbe used to monitor traffic on the taxiways and runways of an airport. It is effective in remote areas orin mountainous terrain where there is no radar coverage, or where radar coverage is limited. One ofthe biggest advantages of ADS-B is the fact that it provides data to pilots and ground controllers, sothat for the first time, they can both view the same information.

99. Once an aircraft has generated an ADS-B transmission other aircraft and ground stationswithin about 100 miles receive the data link broadcasts and display the information in user friendlyformat on a computer screen. Pilots in the cockpit see the traffic on a Cockpit Display of TrafficInformation (CDTI). ADS-B data can be seen by controllers on the ground on their regular ATCdisplay along with other radar targets.

100. A version of ADS-B is currently in operation in Europe and a similar system will beoperational in the United States in 1999. By late 1999, in the second phase of ADS-Bimplementation, software will be developed to provide additional functionality. This software, calledConflict-Detection and Resolution will provide visual and audible cues to pilots when it determinesthat there is the possibility that another aircraft is on a collision course or will pass too closely. Thesystem will be capable of making this assessment at greater distances than are possible with currentTCAS equipment.





101. Since ADS-B uses relatively simple digital technology, it can be scaled down for use in smaller,general aviation aircraft. ADS-B provides an opportunity for small single engine or twin engineaircraft to have cockpit displays similar to the ones in commercial airliners. Accessibility to smalleraircraft therefore gives ADS-B the potential to dramatically improve aviation safety.

Applications of GPS - Area Navigation (RNAV) Approach Procedures102. The following paragraphs, which are extracted from ICAO Doc 8168-OPS/611 Vol 1, givethe guidelines for using basic GPS for RNAV approaches.

Pre-Flight Procedures103. All GPS IFR operations should be conducted in accordance with the aircraft operatingmanual. Prior to an IFR flight using GPS, the operator should ensure that the GPS equipment and theinstallation are approved and certified for the intended IFR operation. The equipment should beoperated in accordance with the provisions of the applicable aircraft operating manual. All pilots/operators must be thoroughly familiar with the GPS equipment installed in the aircraft and itslimitations.

104. The pilot/operator should follow the specific start-up and self-test procedures for the GPSreceiver as outlined in the aircraft operating manual. Basic GPS receivers are capable of non-precision approach operations only.

105. Prior to any GPS IFR operation, a review of all the NOTAMs appropriate to the satelliteconstellation should be undertaken.





106. The pilot must select the appropriate airport(s), runway/approach procedure and initialapproach fix on the aircraft’s GPS receiver to determine RAIM availability for that approach. Airtraffic services personnel may not provide any information about the operational integrity of thesystem. This is especially important when the aircraft has been ‘cleared for the approach’.Procedures should be established in the event that GPS navigation failures are predicted or occur. Inthese situations, the operator may rely on other instrument procedures.

107. Aircraft which are navigating by GPS are considered to be RNAV-equipped aircraft, and theappropriate equipment suffix should be included in the flight plan. If the GPS avionics becomesinoperative, the pilot should immediately advise ATC and amend the equipment suffix forsubsequent flight plans.

GPS Approach Procedures108. GPS receivers must include integrity monitoring routines and be capable of turn anticipation.

109. The airborne navigation database must contain all way-points for the published non-precision approaches to be flown and for the current AIRAC cycle. To ensure the correctness of theGPS database display, pilots should check the data displayed as reasonable for the GPS approachafter loading the procedure into the active flight plan and prior to flying the procedure. Some GPSreceivers provide a moving map display which aids the pilot in conducting this reasonableness check.

110. The approach cannot be flown unless that instrument approach is retrievable from theavionics database.

111. The GPS avionics must store all way-points depicted in the approach to be flown and presentthem in the same sequence as the published non-precision instrument approach procedure chart.





112. Approaches must be flown in accordance with the aircraft operating manual and theprocedure depicted on an appropriate instrument approach chart.

113. Operators must be intimately familiar with their State’s GPS implementation procedures.Some States require an IFR alternate airport to have an approved instrument approach procedure,other than GPS or LORAN-C, which is anticipated to be operational at the estimated arrival time.The aircraft should have the appropriate avionics installed and operational to receive the navigationaids. The operator is responsible for checking NOTAMs to determine the operational status of thealternate airport navigational aids.

114. Procedures should be established in the event that GPS failures occur. In these situations, theoperator may rely on other instrument procedures.

115. Some GNSS receivers may provide altitude information. However, the pilot must complywith the published minimum altitudes using the barometric altimeter.

NOTE:

Now read Chapter 12




FIGURE 209



062 radio navigation (jaa atpl theory)

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