INTEGRATED SYSTEM FOR AUTOMATIC LANDING USING DIFFERENTIAL GPS

AND INERTIAL MEASUREMENT UNIT

Oipl.-Ing. Thomas Jacob

Institute for Guidance and Control of o. Prof. Or.-Ing. G. Schanzer

Technical University of Braunschweig Hans-Sommer Str. 66

0-3300 Braunschweig. West Germany

Phone 0531-391-3716, Telefax 0531-391-4587 Telex 952526tubsw

405 K.-P. Schwarz et al. (eds.), Kinematic Systems in Geodesy, Surveying, and Remote Sensing© Springer-Verlag New York Inc. 1991

ABSTRACT

GPS-receivers have accuracy problems in high precIsion flight gui­dance applications. In dynamic flight maneuvers they show not only operational problems due to satellite masking but also a reduction in accuracy in accelerated flight and turn flight.

In the presented "Integrated System" those problems are solved by integrating GPS in differential mode with inertial measurment sensors into a hybrid system. This integrated system computes a high precise position, flight path and attitude information of a moving platform e.g. an aircraft.

The error behaviour in stationary and in dynamic applications is ex­plained. From the error behaviour a system concept of a hybrid Inte­grated Flight Guidance System is derived.

The position information, estimated in real-time, is used for a flight guidance value generator. These informations are fed to a flight di­rector instrument in the cockpit, which is used by the pilot for manual flight or is fed to an autopilot for automatic flight including automatic approach and touch down.

The system fulfills extreme accuracy requirements and can be used in approach and landing up to ICAO (International Civil Aviation Organi­zation) CAT III. It allows to perform landings even in bad weather conditions. As the integrated system is space based, it computes a landing aid which allows landing at any airfield, not equipped with con­ventional Instrument landing System or Microwave landing System.

In July 1989 the worldwide first automatic landing, using the presen­ted system, based on GPS has been performed by the Institute for Guidance and Control of the Technical University of Braunschweig.

The suitability of the concept (Kalman Filter coupling GPS and inertial measurement units (lMU», for flight path guidance and the accuracy of position finding (better than 1.3 m) will be presented by means of flight tests in a commuter aircraft (OORNIER DO 128) and also by si­mulator results. An inflight comparison of a reference Instrument landing System with the Integrated System shows the accuracy.

INTRODUCTION

When the American Global Positioning System is fully operational, there will be for the first time a navigation system available, which has a higher accuracy than any actual used long, medium or short ran­ge navigation system (except landing systems>. If it is possible, as shown in this paper, to improve position accuracy of GPS CIA code receivers in real-time applications to the order of 3 m or less, even an approach of an aircraft could be performed by using a space based "Integrated Navigation System". When GPS is available world­wide, approach and landing could be performed at any place not equipped

406

with conventional Instrument Landing Systems (ILS) or Microwave Lan­ding Systems (MLS), even in bad weather conditions. This can be advantageous for General Aviation and for feeder services, which normally operate from small airfields to international airports. Especially at small airfields no landings are allowed in bad weather conditions, because these fields in general are not equipped with ILS or MLS. Also most airports in the third world are not equipped with ILS or are equipped with systems which may not be reliable due to insufficient maintenance. Using the highly accurate Integrated Navigation System flight safety could be increased significantly at these fields.

The "Integrated Navigation System" as it is presented in this paper allows from a technical point of view to gUide an aircraft, by using ane system, from any terminal A to any terminal B including:

- runway - takeoff - approach - landing

guidance.

In developping an Integrated Guidance System for landing approach guidance, the following aspects:

- accuracy - dyn am ice r r or char ac t er i s tics - integrity

must be considered for the complete flight guidance system.

ACCURACY REQUIREMENTS FOR INSTRUMENT LANDING SYSTEMS

The accuracy and performance requirements for the ground equipment of Instrument Landing Systems (lLS) have been defined by the Inter­n at ion a I C i v iI A v i a t ion 0 r g ani sat ion (I C A 0) inA nne x 1 0 I 1 I. The r e qui -rements differ for horizontal and vertical guidance instruments. For each, a maximum bias value has been defined. That is an angular shift of the mean nominal path and a maximum beam bend, which is mainly due to multipath. The values are defined depending on the visibility conditions. For a horizontal visibility of 200 m and a vertical visibility of 0 m the ILS ground equipment has to fulfil the requirements for ca­tegory CAT III (The following values are calculated for a standard runway with a 3 degree glide path and 3000m runway length for CAT III ):

407

offset bend

horizontal [m] 3.0 2.4

vertical [m] 0.6 0.5

Tab!.@. 1~ ac curacy req uirement s for IlS ground equ ipment

The requirements for the guide beam characteristic are set in such a way. that the deviation of the aircraft due to errors of ground instru­ments at the runway threshold are less than:

CAT

horizontal vertical

III

5.0 m 1.2 m

Additionally. the roll and pitch attitude should not deviate more than 2 degrees from the reference values at the threshold due to course bends. For CAT III the signal quality has to be good enough to provide automatic flight.

If the Integrated Flight Guidance System should be applied in a com­mercial airplane for landings. it must satisfy the requirements of the ICAO depending on the weather conditions. Although GPS in the presented system is just one of many different sensors it is this sensor. that is responsible for stationary accuracy. Therefore. the error characteristics of a high precision 5 channel GPS C/ A code receiver in stationary as well as in dynamic flight tests have been analysed.

ERROR CHARACTERISTICS OF (iPS RECEIVERS IN REALTIME DIFFERENTIAL MODE

The typical error behaviour of GPS-measurements in stationary appli­cations using only block I satellites without selective availability (SV: 3.9.11,12.13) is shown in figure 1.

408

Nord­position

1m)

3

2

57'00 57600 57800 58000 58200 58'00 58600 58800 59000 tis)

Fig'yr.~ 1l Measurement of nort position using 5 channel GP5-CI A-code r e c e i v era n d b I 0 c k I sat e II i t e s wit ho u t 5 I A

These measurements have been made using the raw data of a 5 chan­nel CIA-code receiver, calculating a complementary filter of carrier phase and code measurements. This error behaviour in stationary flight can be modelled in the following way:

offset 15m drift 0.1 mlmin

two oscillations with

amplitude A1 0.2 m, period T1 16 s amplitude A2 0.6 m, period T2 250 s.

88 68 1&8 28 8

-28 -118 -68 -88

-188

North Position [m]

27888 28888 29889

lime [sec.l -

Fig'yr.~ 2.l Measurement of north pos it ion using 5-channal GP5 CI A­code receiver, block I and block II satellites

These natural errors are superposed by imitated errors - the selecti­ve availability (51 A) implemented on block II satellites. These errors are implemented into the system with the aim, to allow non-authorized users a maximum horizontal accuracy of 100 m (2drms) and a vertical

409

accuracy of 170 m. The influence of S/A in a measurement session versus 5 hours is shown on figure 2. In these measurements the same receiver with exactly the same software has been used operating with three satellite configurations: SV:9, 11, 12, 13, 18 and 9, 12, 13, 18, 20, and 3, 9, 12, 13, 20), A plot of north position versus east position of the same measurement is given in figure 3. It can be recognized from these measurements that a configuration with just one block II satellite produces a typical position error of 20 - 30 m, while confi­gurations with two block II satellites show a positon error of up to 140 m.

158· ft==t===jf====t====t===t====t==91

le8 · "--+---II--__l---4I--~.t<==_J--__l1

1.1 II-__l----l -\- =---j---++t---j---+---jl

-58 . II-__l----If--__l-;~..::....qb__~__l----II--__li

- 118· H- -I---I

-21. --' .8 21.1 U . I 61 . ' .... I ....

F i gyr.~ ~.,;. Po sit ion plot u sin g 5 c han n e I G P S - C I A - cod ere c e i v e r, b I 0 c k I and block II satellites (same configuration as in fig. 2)

The typical error behaviour of stand alone GPS can be modelled as follows:

- the measured GPS-position has an offset versus the real position - the measured GPS-position drifts, due to the movement of the

satellites (change of constellation, GDOP and S/A) even when there is no aircraft movement

- typical measurements have a noise, of which the amplitude is de­pendent on receiver quality;

- an oscillation is superposed (receiver dependent, multipath, S/A) - when the selected satellite constellation is changed due to the

rising of a new satellite, the position measurements react nearly like a step function with an amplitude of several meters (errors in the order of 8 m have been measured)

410

As generally known, the offset and the time dependency of the offset can be extensively eliminated by using Differential-GPS-Technique (fig. 4>.

Regional-Airfield

Fig~I.~ !~ 0 ifferent ial GPS for precision approach guidance

As the position of the ground antenna is known, the range from the ground to each satellite can be calculated. By comparing the computed range with the measured range the actual system error can be deter­mined. Transmitting these errors to the aircraft via a telemetry, they can be corrected in the onboard position finding computation if the di­s tan c e tot h e g r 0 u n ds tat ion iss m a II en 0 ugh. T 0 use 0 - G P 5 for an a p­proach and landing guidance system the most interesting questions are:

- is it possible to correct the S/A - effect in real-time to an error level below 4m horizontal?

- which is the maximum distance between ground receiver and air­borne receiver, up to which the accuracy requirements are met.

Some measurements indicate that distances less than 200 km can be adequate. In optimal conditions (low range between user and reference ground station, no multipath. same propagation path satellite- reference sta­tion as satellite-user and no system degradation> this technique can reduce the position error up to the receiver residuals which are in the order of a meter or a few centimeter depending on the receiver type (fig.5>.

411

xgd il t 1.1 I m I -1.2

.... ~ 57111

Y9diftl.71~ Iml 1.11 ~

1.51 ~WIIMU1l UI ~==r.

57~6H~7-snMI~I~5~7a~II~7-57~II~I~5~a"~I'-g~IH~7.5a~2'~I~S~a!~11'-"1M~n=-'5~a5~1I'-<'5""U~<g~7~U~s~a.~'I~ tIs) _

:!~;'Ij:~~~ 57611 577., 578.. 57911 5S... 581'1 582el 583.. 58U' gSiI 58811 g1ll1 5.8 •• tls1 _

Fig'yr.~ 5: Real-time O-GPS using bloc k I satellites (tau =0.6 s, v=O. mIs, 5 channel CIA code, L1 receiver)

In figure 6 results of a real time O-GPS technique, as developed in the Institute for Flight Guidance and Control, are calculated from the same measurements as in figure 3.

North [mJ

to' 4\

-2.' W 1'1

-6 . •

-: ---1.9

-9.9

- II.

- 13. - 13. ,-II. - 9.5 -7.! -6. 1 -... -2.1 ' .1

East [mJ

Fig'yr.~ .§..;, Real-time O-GPS using block I and block II satellites h=0.6 s, V=O. mIs, 5 channel CIA code, L1 receiver)

This has been a mainly stationary measurement versus 5 hours. In this session one antenna has been moved 8 m to the south and 1.5 m to the east. Although this measurements are stationary with an extremly short

412

baseline they show the potential of this real time technique. Although there has been the aim to degrade a CIA-code user by the SI A to a horizontal accuracy of 100m and vertical to 170 m, figure 6 shows, that it is possible

- by using the online real-time D-GPS technique, to correct the sy­stem degradation to

98 X horizontal 96 X vertical

- to fulfil the horizontal requirements of the ICAO for CAT III.

However using this correction technique, the dynamic error characteris­tics are not improved. In stationary tests, as well as in flight trials, tested GPS receivers have proved to be excellently accurate in the long term, however, a lower accuracy in dynamic maneuvers has been detected. Depending on the receiver type receivers show

- dynamic errors in flight phases with longitudinal accelerations as well as in phases with lateral accelerations. The reasons for the dynamic errors are:

- an influence of acceleration on the receiver clock (crystal oscil­lator) .

- an influence of changes in the measurement signals to the output due to transfer function of the code tracking loop (delay lock loop) and phase tracking loop of the carrier signal (costas loop).

- dynamic problems due to receiver integrated software and filter technique using low pass filters to reduce noise 12/.

- time lack due to signal processing up to 0.4s 12/. electromagnetic interferences in airborne applications with air­craft radio. This is a special problem of some two frequency recei­vers, when they are operating with an HF-Fllter for both frequen­cies with a too large bandwidth.

- operational problems in airborne applications. Flying a turn with a bank angle greater than the elevation of a satellite, a masking of the antenna is produced as:

- the hemispherical antenna is focused in such a direction that the locked satellites are undetectable. parts of the aircraft (e.g. wing, body, rudder or props) move in­to the line of sight between satellite and user antenna.

This effect produces always cycle slips in the phase measurement and of course a lock-off of the masked satellite. After reducing the bank angle again. receivers need up to 10 s - 30 s to relock again. In the most successful cases this effect produces an error of only several meters (of about the same order when changing the constellation) or at the worst case a total loss of GPS position information (fig. 7 point A.)

413

4000

i 3000 .-~

.---... -------......... '\ e n \\ #

oJ: 2000 A .... ~ lv 0 z

1000

(\( 0 --- ? -1000 -4000 -2000 0 . 2000. 4000

East m ~

Fi.9!!.!.~ I~ Ground Track measured with real-time O-GPS using block I satellites in a flight test <'r=0.6 s, V=60. mis, 5 channel CIA code, L1 receiver, 360 0 circ with bank angle at point A <I> =60 0 , at B standard curve <I> =20 0 )

The O-GPS measurements in a flight test are shown in figure 7. The pilot had the task to fly three flight patterns. In one flight pattern he had t 0 fly a 3 6 0 0 c i r c lew i t h a ban k an g leo f u p t 0 6 0 0 nor tho f the airport. In this extreme maneuver three satellites were masked. There­fore the GPS receiver was not able to calculate any position, not only during the maneuver but also for the next 45 sec. as he had to relock again. Therefore for several kilometers the pilot had no position infor­mation from GPS.

For a high precision flight guidance during a landing phase even in CAT I conditions none of these errors can be accepted. Altogether the errors must be for CAT III at least less than 3.0 m in the horizontal and 0.6 m in the vertical. In addition to the accuracy the ICAO re­quires integrity for the landing guidance system. In case of a malfunc­tion of a CAT IlLS glidepath transmitter the pilot has to have an "off­signal"· within 6 sec. All these requirements are met in the Integrated Flight Guidance System by system integration of complementary sen­sors (GPS and INS), including adequate filtering.

CONCEPT OF THE INTEGRATED FLIGHT GUIDANCE SYSTEM

To improve the dynamic behaviour of the system and also for safety reasons during a landing approach it is a need to generate additional position information without using GPS.

In dynamic maneuvers the aircraft position, (cp, A, h), the inertial ve­locity V, the attitude angles <1>, e, tV (bank, pitch, azimuth) and the bo-

414

dy fixed accelerations, can be measured and calculated by using an inertial measurement unit (IMU). IMU's have a good short term accu­racy, however. in the long term they have recognizable drifts. The long te-rm accuracy is dependent on the gyro drift which determines the coordinate axes misalignement En. Ea. Ed' With today high quality IMUs, a typical drift of 0.5 m/s can be obtained. Obviously a system concept which utilizes the good long term accuracy of the GPS and the good short term accuracy of an IMU would produce a good overall accu­racy. This concept is realized in the "Integrated Flight Guidance System" (fig. 8),

Pilot

CaImands

F i .9.Y.!:.~ ~.!. I n t e g rat e d F I i g h t G u ida n c e S y s t e m

The "Integrated Flight Guidance System" is basically composed of two parts:

- a position finding system and - a guidance generator.

To improve the dynamic characteristics of the entire system and to get sufficient information concerning the flight path during a breakdown of satellite information. the integrated flight gUidance system is coupled by Kalman filter technique with inertial sensors - gyros and accelerome­ters. Radio or barometric altitude sensors are used additionally to 0-GPS if a satellite masking is produced. In the Kalman filter there are im­plemented error models, that estimate the errors of system states as well as the errors of the sensor systems, as there are:

navigation coordinate misalignement gyrodrifts offsets of the accelerometers receiver clock error receiver clock error drift

415

By using these models the filter learns the errors of the inertial sen­sors in those phases of flight, where the GPS-receiver is in a good con­dition. If some satellites are masked in maneuver flight, the hybrid po­sitioning system calculates a position information using online calibra­ted inertial sensors. Therefore it is possible to use low cost inertial sensors instead of expensive inertial platform systems, and to calcu­late a position with high accuracy. To the output the system gives on­line calibrated

- position - velocity - acceleration - attitude angles - attitude angular rates

with an update rate of up to 50 Hz.

Although the best estimation of position is calculated using this tech­nique, no pilot is able to follow a nominal flight path using only a position information in latitude, longitude and height. It was advisible to display the position information to the pilot in a conventional manner. Therefore, a nominal flight path consisting of standard rate turns and linear parts using known coordinates of the target place (waypoint, airport threshold) is calculated. By calculating the nominal flight path s"tarting from the precisely known actual position a substantial advantage is achieved be­cause no intercept maneuver has to be flown.

Flight guidance data are calculated from the information of the actual position, the deviation and the attitude relative to the nominal flight path. With this procedure the pilot receives information on how to fol­low the nominal flight path, which may be curved in the horizontal as well as in the vertical plane and how to recover if he deviates from it. For the indication during the flight test an ILS cross deviation indica­tor or a flight director can be used.

SIMULATION AND FLIGHT TEST RESULTS

Fig. 9 presents simulation results. In the simulator the same flight procedure has been flown as in the flight test in fig. 5. The position errors of the subsystem GPS and INS (1 NM/h) as well as the errors of the integrated system are plotted versus time.

416

t 70

! 60 Xi ~ 50

40

30

30

10

0

0 60 120 180 240 300 360 t (sl-

Position error of INS (1Nm/h). GPS and Integrated Flight Guidance System (t=150 s ~=600. t=241s ~=200)

At t=140 sec a circle with a bank angle of 60 0 has been flown. While flying 60 0 bank angle a masking of 3 satellites has been produced for 100 sec. In this time GPS gives no position information. While the INS has produced a position error of 48 m. beginning at the alignment to the end of the masking. the error of the integrated system is 71 X (14 m) lower than the error of the INS. At t=260 sec the aircraft flew a turn returning to the extended center line (ECl)' In this phase of the flight. the satellites were in a bad constellation (Geometric oelution of Precision GoOP > 10). One satellite was masked. Here the error of the GPS grows up to 50 meters. The integrated system. however. has produced an error of up to 7m which is 90 X lower than the error of the Inertial Navigation System. although there had been no GPS posi­tion measurement for 65 s!

If there would be a loss of satellites during an approach the integra­ted system is able to recognize the satellite error. The system checks the errors of the satellites on ground. realizes failures. reduces their effect and gives a signal to the pilot. Considering the time until leaving CAT II accuracy requirements the pilot would have up to 40 5 to conti­nue the approach. to wait for relocking the satellites or to make his de­cision for a go around maneuver. Especially in this point the Integrated System has significant advantages compared to conventional IlS or MlS. It has a high integrity!

These simulations show. that by combining IMU and GPS in the discussed integrated system an overall accuracy can be achieved that is up to 90 X better than the accuracy of an INS standing alone. even in phases where no GPS is available. In phases where GPS is available the accuracy of alteration of position is in the dcm - order. using differential GPS phase measurements. however with an performance during high dynamic flight

417

phases which is much better than GPS standing alone.

4000

! 3000

~ 2000 '­o z:

1000

o

-1000~--~_4~0~00~--~~----~------~----~---- 2000 0 2000 4000

East m _

Figy'r.~ lQ..;. Ground Track computed by the "Integrated Flight Guidance System" (same flight test as in fig. 7)

Fig. 10 shows the ground track of the same flight as that displayed in fig. 7. The GPS-position and the position output of the position-fin­ding part of the "Integrated Flight Guidance System" are presented here. Based on the sensor errors, which are determined by the system, the position-finding part is able to determine the flight path and the attitude angles with a high precision. Even with a breakdown of the GPS signals, the filter algorithms still give the position, speed and attitude angle with a high precision for a limited time. A decisive prerequisite for that is a realistic mathematical model of the dynamic error charac­teristics of the inertial systems.

Which overall position accuracy can be achieved using the "Integrated Flight Guidance System"? The flight test results in automatic approach are shown for the lateral and vertical position in fig. 11. It can be recognized from these plots that an autopilot is able to fly the aircraft in a stable condition to the runway threshold using only the signals computed by the integrated navigation system. It should be emphasized that by using this integra­ted system the procedure of flying an intercept maneuver is not ne­cessary, because the aircraft is always under cover of the navigation system. Additionally this system is able to fly not only an automatic a p pro a c h but a Iso a n aut 0 mat i c to u c h dow n (i n c Iud i og aut 0 t h r 0 tt Ie) .

Figure 12 shows a comparison of the computed deviation from the glide path calculated by the "Integrated Flight Guidance System" with the measured Instrument Landing System of runway 27 R in Hannover (which is regarded by the ICAO as a reference system in DOC 8071 and meets the ICAO requirements of Annex 10 for CAT III ILS)' In se­veral approaches it could be verified that the difference of both glide path deviations (horizontal fig.12a and vertical fig. 12b) is within the requirements.

418

If the ILS is regarded as a reference system without any error the difference signal would be the total error of the integrated system. In reality, however, there is an error contribution by the ILS, so that the accuracy of the Integrated system is better than the difference sig­nal.

Artificial Glide Path

200 F1i;rtPath

0 0 1000 5000.

Distanct to touch down point [mJ -

4000 North Position Iml

3000

2000

1000

oL---------'----r Extended Centerline

o 1000 2000 3000 4000 5000 6000 East Position 1m)

Fi.Q.YI.~ 11.;. Ground Track and vertical flight path computed by the "In­tegrated Flight Guidance System" in automatic approach

419

I .... '.35

~'.3' C '.25

~ '.2' :i: '.15

~ '.1' ;: ' •• 5 ca .... t--+---4-----, ....... -I--+-....,.,.-+-Ir---I-=-'I-~~~~ j-•. '5 -'.1' -'.15

-'.2' -'.25

-'.3' -'.35 - ..... 1::.-.I_---,.--L __ -L-~_...I_ __ .L-_--lL..-.-'-__L_.l_._L 12... 1.... _. . ... 2... • -2'"

D •• Thr " -

GPOIf'-GPS-ILS. GPDlrr-GPS-Radioheight 1.21 • 1.18 1.18

dI 1.'" Ir-------A--kl ~ 1.12 ':: 1.11 e '.18 !! '.18 a ;;: .....

' •• 2

I CAO Annex 10 CAT. req. ,./"

.... ~---~~-~~~--~~~+-~-~~ -1.12 -1.1" -1.1& -1.18 -1.1' -1.12 -1.'" ......w'--------....---1.18 -1.18 -•. 2' ~~~~~-~~---'""""""'~ ___ -L-__ -'-.--I'----L

12'" 1 .. '. a". 68" ... ,. 2.'. • DleThr

Fi.9.Yr.~ 1~~~1~.b.1 Comparison of ILS and Integrated Navigation System with ICAO re­quirements (a: ~LOCILS-~LOCGPS'

b: ~GPILS-~GPGPS and tJ. G P H r a d i 0 - ~ G P G P S)

420

SUMMARY

The flight tests which have been made with the "Integrated Flight Guidance System", developed at the Institute for Guidance and Control of the Technical University of Braunschweig have demonstrated good results by combining two sensor systems with different, complementary, time dependent, signal qualities: the inertial sensors, with their excellent short-term characteristics; and the GPS with excellent long-term c h a r act e r i s tic s. Wit h the K a I man fil t e r t e c h n i que it is po s sib Ie, eve n in high dynamic flight phases, to determine a position of high precision and reliability. The position determined in flight tests is better, than the precision of each system standing alone. While in real-time application the vertical accuracy of the "Integrated Flight Guidance System" is sufficient for CAT II landings the horizontal accuracy fulfills the de­mands of the ICAO for CAT III requirements.

It is possible to use the presented system in all high precision naviga­tion applications.

In offline applications of the position finding part of the "Integrated Flight Guidance System" accuracies in the order of 10cm can be achieved, by using special differential techniques (double differenti­al),

The first application of differential GPSIINS guidance will probably be the use as a high preCision position reference system and may be in regional air traffic the application as a landing gUidance system for such airports which have no IlS or MlS systems installed. It may be expected that a discussion about whether

- the Microwave landing System (MlS) might be out of date before it is generally used

- the Global Positioning System (GPS) can be used instead of MlS especially for business- or commuter aircraft operating into small airfields

will arise in the near future.

421

BIBLIOGRAPHY

/1/ Aeronautical Telecommunications, International Civil Aviation Organisation (ICAO) ANNEX 10

/2/ Avionics Navigation Systems Kay ton, M.; Fried, W.; John Wiley & Sons 1969

/3/ Richtlinien fur den Allwetterflugbetrieb nach Betriebsstufe " N a c h ric h ten fu r L u f tf a h r erN f L 1- 3 5 0 / 7 2 Bun des a n s t a I t fur Flugsicherung

/4/ Principle of Operation of NAVSTAR and System-characteristics Milliken, R.J; Zoller, C.J. AGARDograph AG 245

/5/ The NAVSTAR GPS System AGARD Lecture Series No. 161

/6/ Differential Operation of NAVSTAR GPS Kalafus, R.M. Journal of the Institute of Navigation Vol. 30, 1983

/7/ Flight Guidance" G. Schanzer, Technical University Braunschweig

/8/ Techniques of the development of Error Models for Aided Strapdown Navigation Systems Lechner, W. AGARDograph AG 256

/9/ Genauigkeitsanalyse von Tragheitsnavigationssystemen N. Lohl, Technical University Braunschweig 1981

/10/ Integrated flight guidance system using Differential-GPS for landing approach guidance Th. Jacob, AGARD Guidance and Control Panel Lissabon 1989

/11/ Approach flight guidance of a regional air traffic aircraft using GPS in differential mode Th. Jacob, ICAS Jerusalem 1988

422