US005923286A
`5,923,286
`[11] Patent Number:
`(15
`United States Patent
`Divakaruni
`[45] Date of Patent:
`Jul, 13, 1999
`
`
`[54] GPS/IRS GLOBAL POSITION
`DETERMINATION METHOD AND
`APPARATUS WITH INTEGRITY LOSS
`PROVISIONS
`
`[75]
`
`Inventor: Sudhakar P. Divakaruni, Scottsdale,
`Ariz.
`
`[73] Assignee: Honeywell Inc., Minneapolis, Minn.
`
`[21] Appl. No.: 08/735,764
`
`[22]
`
`Filed:
`
`Oct. 23, 1996
`
`Int. CL.° ...
`[51]
`. GOIS 5/02; HO4B 7/185
`
`
`[52] US. C1 ee eeceeeetieceseseseecenseenesnees 342/357; 701/213
`[58] Field of Search oe 342/357, 450,
`342/457; 701/213, 214
`
`[56]
`
`References Cited
`U.S. PATENT DOCUMENTS
`
`
`
`2/1989 Harral et al. wo... eee eeseeee eee 342/451
`4,806,940
`2/1995 Kao oe
`.. 364/450
`§,394,333
`5,461,388 10/1995 Applegateet al.
`wee 342/357
`5,504,482
`4/1996 Schreder.......
`.. 340/995
`5,504,492
`4/1996 Class et alo ccc cssece ese 342/357
`5,512,903
`4/1996 Schmidke oes 312/357
`8/1996 Buchleretal.
`wee 342/357
`5,543,804
`
`we 364/443
`5,583,774 12/1996 Diesel
`.......
`
`. 364/449.1
`5,606,506
`2/1997 Kyrtsos.....
`
`8/1997 Ebneretal. ..
`wees 342/357
`5,657,025
`FOREIGN PATENT DOCUMENTS
`
`0 629 877 12/1994 European Pat. Off.
`95 34850 12/1995 WIPO.
`
`.
`
`OTHER PUBLICATIONS
`
`Article entitled Minimum Operational Performance Stan-
`dards for Airborne Supplemental Navigation Equipment
`Using Global Positioning System (GPS), document No.
`RTCA/DO-208,Jul. 1991, prepared by: SC-159, pp. 19-22.
`Appendix F entitled “Least-Squares Residuals RAIM
`Method” from document No. RTCA/DO-208, Jul. 1991,
`prepared by SC—159, pp. 1-4.
`Article entitled “Implementation of a RAIM Monitor and a
`GPSReceiver and an Integrated GPS/IRS”by Mats Brenner,
`in the proceedings of TON GPS-90, Third International
`Technical Meeting of the Satellite Divisionofthe Institute of
`Navigation, Sep. 19-21, 1990, located at p. 397.
`Patent Abstracts of Japan, vol. 013, No. 135 (p-851), Apr. 5,
`1989.
`
`Primary Examiner—Thomas Tareza
`Assistant Examiner—Dao L. Pham
`
`Attorney, Agent, or Firm—Charles J. Ungemach; Ronald E.
`Champion
`
`[57]
`
`ABSTRACT
`
`A system for use with an inertial reference system and a
`global position receiver for calculating a position error after
`a loss of integrity by utilizing the global position system
`values for position and velocity at a time just before the loss
`of integrity and by utilizing the inertial reference system
`position modified by the known error in inertial reference
`system position as it varies with time and the position error
`as calculated by the global position system velocity extrapo-
`lated over the time since integrity loss.
`
`19 Claims, 3 Drawing Sheets
`
`156
`
`I1
`
`
`
`120
`
` GPS SATELLITE
`SIGNAL
`
`
`
`185
`RECEIVER
`
`
`
` 150 152
`
`
`
`RAIM
`
`
`HIL/VIL
`INTEGRITY
`
`LIMIT COMPARATOR
`GPS
`
`
`POSITION /SOLUTION
`
`
`INFORMATION
`MODIFIED
`Vv
`
`
`IRS POSITION
`
`
`
`
`[|
`
`GPS INTEGRITY
`VALIDATION
`NIT
`ASN
`
`PROCESSOR
`
`II'1|1I1I't|(!ll||'
`
`
`
`a
`
`INERTIAL
`SENSORS
`
`160
`
`Google Exhibit 1021
`Google Exhibit 1021
`Google v. Mullen
`Google v. Mullen
`
`cI
`POSITION
`
`DIFFERENCE
`
`
`CALCULATOR
`
`240
`
`
`REFERENCE [
`
`
`ALTERNATE
`r
`POSITION
`
`
`INTEGRITY
`1
`STIMATOR
`DATA_IRS
`E
`
`LIMIT COMPARATOR|
`REFERENCE TT
`
`IRS
`|
`
`
`POSITION JANERTIAL
`INFORMATION
`
`
`PROCESSOR
`
`

`

`U.S. Patent
`
`Jul. 13, 1999
`
`Sheet 1 of 3
`
`5,923,286
`
`Sst
`
`I|
`
`YOLVuVdNOD
`
`LINALIMDALN|Ty
`
`
`WIVIH
`
`Oly
`
`CZ
`
`UyxOldd)
`
`OSl
`
`yOSS4008d
`
`WIVSdd
`
`ZZ
`
`08|zt|
`
`j|
`
`
`
`;\LOlldALIMOSLNISAO
`||(18a=NOLLWOMVWA
`
`SNVIN
`
`cCl
`
`cel
`
`cel
`
`NOILMIOS/NOILISOd
`
`NOILVWHO4INI
`
`NOSS300Nd
`
`eral
`
`SALIMNSLVS
`
`WNOIS
`
`YSAISO4Y
`
`LAS
`
`CAS
`
`
`
` 8lI-——|O|OlavyvVLVva
`
`
`
` YOLVANILSS!vilOltNOILISOd|yu3"xXSOd
`
`weaASOd
`
`SdVLVG
`
`cel
`
`Surviva
`
`orl
`
`JONLNOD
`
`LHONAGAHSOd
`
`|
`
`||
`
`091
`
`!Qik
`
`ZL
`
`ce
`
`Sul
`
`WILMSNI/NOILISOd
`
`yOSS490dd
`
`NOILVANYOSINI
`
`TWILYSNI
`
`SYOSN3S
`
`
`
`
`
`
`

`

`U.S. Patent
`
`Jul. 13, 1999
`
`Sheet 2 of 3
`
`5,923,286
`
`II|I|I|I|lIII||I||I||
`
`
`
`O¢¢222
`
`JS0N3435434
`
`ALINOSLNI;3LVNYZLIV
`
`LVYVdWODLIN
`
`YOLVAILSS
`
`ZLL
`
`oT4crenowisod|°%vel
`
`!GAHSOdNOLLWAYOINI1prapvWILYINI/NOLLISOd
`IJONSHqIIN=|YO|NOILISOd
`
`0zzILMIOS/NOILISOd
`|2bISul
`
`1aS~SOdSYOSNIS“4OLOSTaSoFWILYANI
`9stZSIOst
`
`ccU3AI3034
`
`NOLVUVGNODLINMYOSS3908d“osc
`JOULNODcee
`|YN
`NOILISOdSHUIALINDALNISd9||[NOSSIOONd
`
`
`Q3i4!dOWOZtazNOLLVWYOINI
`
`
`ALINOALNIIVYSdOraAl
`
`SS01~sul7sodyOSS3D0"Nd
`ZblEBoztLI
`
`NIVWASH
`
`SLIVISLVYSSd9
`
`WNOIS
`
`
`
`SSO1~SHI”~SOd|YOLVWILSS
`
`OvZa
`
`NOILISOd
`
`JONI3435410
`
`
`
`YOLVINDIWS
`
`NOILVOIIVA
`
`NI?AI
`
`youeTvivd
`NOLLVYSTIOOV/ALIOOTSA
`
`Leh
`
`cee
`
`SdVLVG
`
`yOUNS
`
`NOLVWILSS
`
`Old
`
`
`
`
`
`
`
`
`
`
`

`

`Jul. 13, 1999
`
`Sheet 3 of 3
`
`5,923,286
`
`U.S. Patent
`
`oo¢
`
`Yaya~X™135A
`
`yOYNSVLVd
`
`cee
`
`Sul
`
`X T
`
`SA
`
`

`

`5,923,256
`
`1
`GPS/IRS GLOBAL POSITION
`DETERMINATION METHOD AND
`APPARATUS WITH INTEGRITY LOSS
`PROVISIONS
`
`BACKGROUND OF THE INVENTION
`
`1. Field of the Invention
`
`The present invention pertains to a system employed for
`determining the global position of a mobile unit by employ-
`ment of both an inertial reference system (IRS) and a
`satellite positioning system (GPS), and more specifically, a
`system which employs a provision for determining the
`mobile unit’s global position and corresponding integrity
`during those periods of time in which the GPSsatellite
`constellation is insufficient for establishing GPS integrity
`limit values by employment of RAIM.
`2. Description of the Related Art
`Satellite positioning systems are now well-knownin the
`art. Such systems, for example, NAVS'TAR-GPS,are rapidly
`being employed for a determination of the geocentric posi-
`tion of mobile units, such as water and land vehicles, space
`and aircraft, and survey equipment, to name a few.
`In aircraft, GPS systemsare being utilized for navigation,
`flight control, and airspace control. These GPS systems may
`operate independently or in combination with inertial refer-
`ence systemsorattitude heading reference systems in order
`to provide information particularly during a flight mission.
`Global positioning systems, hereinafter referred to as
`“GPS”, similar to NAVSTAR, commonly use a GPS
`receiver, located on a mobile unit, for receiving satellite
`information signals transmitted from a pluralityof satellites.
`Each GPSsatellite transmits a satellite information signal
`containing data that allows a user to determine the range or
`distance between selected GPS satellites and the antenna
`associated with the mobile unit’s GPS receiver. These dis-
`
`tances are then used to compute the geocentric position
`coordinates of the receiver unit using known triangulation
`techniques. The computed geocentric position coordinates
`may, in turn, be translated to earth latitude and longitude
`coordinates.
`
`In order to determine the position of the GPS receiver, a
`minimum of four unique satellite information signals are
`required, rather than the expected three (three position,
`unknown coordinates). This is so, since the GPS receiver
`generally includes a receiver clock whichis not as accurate
`as the atomic clock normally associated with each of the
`satellites. Therefore, receiving satellite information signals
`from four different satellites provides a complete solution
`which permits the correction of any receiver clock error as
`is well-understood in the art. [lerein,
`the GPS receiver
`position derived bythe triangulation technique using data
`from multiple satellites is referred to as the “GPS estimated
`position”,
`identificd as POS_GPS. The accuracy of this
`estimated GPS position is dependent upon many factors,
`including, among others, atmospheric conditions, selective
`satellite availability, and the relevant position of the satel-
`lites with respect to the line of sight view ofthe satellites.
`Associated with a GPS estimated position is a “position
`error bound” as particularly defined by accepted GPSsys-
`tems standards which have been developed by the Radio
`Technical Commission for Aeronautics (RTCA), in associa-
`tion with aeronautical organizations of the United States
`from both governmentand industry. The RTCAhasdefined
`the phrase “GPS system integrity” as the ability of a GPS
`system to provide timely warnings to users when the GPS
`
`10
`
`15
`
`30
`
`35
`
`40
`
`45
`
`50
`
`55
`
`60
`
`65
`
`2
`system should not be used for navigation. “Systemintegrity”
`is particularly identified in a documententitled “Minimum
`Operational Performance Standards for Airborne Supple-
`mental Navigation Equipment Using Global Positioning
`System (GPS)”, document oumber RTCA/DO-208, July
`1991, prepared by: SC-159, beginning at section 1.5. As
`described therein, GPS is complicated in that it is a four-
`dimensional system involving three components of position
`and one time component. As also described in the aforesaid
`RTCApublication, the signal-in-space error transtorms into
`a horizontal position error via a relatively complex function
`of a satellite constellation geometry at any given moment.
`‘The GPSintegrity system must interpret the information it
`has about the received GPS signals and error terms in terms
`of the induced horizontal position error, commonlyreferred
`to as the “position error bound”, and then make a decision
`as to whether the position error bound is outside the allow-
`able radial error, spccificd for a particular phasc ofthe flight
`mission in progress. The allowable error is referred to as the
`“alarm limit”, herein referred to as the “integrity alarm
`limit”. If the horizontal position error bound is found to
`exceed the integrity alarm limit, a timely warning must be
`issued by the GPS receiver or subsystem to notify the pilot
`that the GPS estimated position should not be relied upon.
`Two rather distinct methods of assuring GPS integrity
`have evolved as civilian use of GPS has progressed. One is
`the Receiver Autonomous Integrity Monitoring (RAIM)
`concept, and the other is the ground monitoring approach
`that goes under the “GPS Integrity Channel” (GIC). The
`intent of both of these methods is the calculation of the
`
`position error bound with regard to the current GPS esti-
`mated position so that it may be compared with the alarm
`limit associated with a particular phase of a flight mission.
`The receiver autonomous integrity monitoring system
`(RAIM) employs a self-consistency check among the
`measurements, more specifically, GPS pseudo range mea-
`surements. Satellite redundancy is required to perform a
`self-consistency check on an instantaneousbasis. Thus,five
`satellites must be in view, ie., five satellite information
`signals received and pseudo range measurements calculated
`by a GPSreceiver. If fewer thanfive satellites are in view,
`the value of the predicted position error bound will be
`infinite. Also, constraints are placed on the satellite constel-
`lation geometry that must be met if the self-consistency
`check is to be effective in the presenceofnoise, e.g., azimuth
`angle of the satellite relative to user position. Generally, a
`satellite constellation with manysatellites in view permits a
`robust integrity monitoring system. Conversely, a satellite
`constellation having only a fewsatellites in view, may limit
`the availability of an integrity monitoring system. ‘hus,
`there may be short periods when a good consistency check
`is not possible (less than five satellites in view). The main
`feature of RAIM isthat it is completely self-contained and
`relatively easy to implementin software.
`Examples of RAIM may be found in the aforementioned
`RTCApublication, Appendix F,andalso in anarticle entitled
`“Implementation of a RAIM Monitor and a GPS Receiver
`and an Integrated GPS/IRS” by Mats Brenner, located at
`page 397, in the proceedings of ION GPS-90, Third Inter-
`national Technical Meeting of the Satellite Division of the
`Institute of Navigation, Sep. 19-21, 1990.
`GPSsystems which incorporate RAIM output a position
`error bound value which represents the probabilistic radial
`errors of the navigation solution, namely, the GPS estimated
`position of the receiver unit. Currently, RAIM maygenerate
`several numbers,
`including, a horizontal position error
`bound value (sometimes referred to as HIL—Horizontal
`
`

`

`5,923,256
`
`3
`Integrity Limit), a vertical position error bound value
`(sometimes referred to as VIL—Vertical Integrity Limit),
`and spherical position error bound for the currenttime,i.e.,
`the instance of time that GPS measurements occurred.
`
`4
`a Global Position System Receiver to provide a second and
`independent source of information about the aircraft posi-
`tion. The two sets of information are mathematically com-
`bined in a Flight Management System (FMS)to determine
`a hybrid position POS_HYB. In turn, this position value
`Once calculated, the position error bound value(s), HIT.
`along with attitude and rate signals from the Inertial Refer-
`and/or VIL, may be compared with selectable integrity
`ence Unit may be providedinaflight control for controlling
`alarm limit values to determine if the pilot can rely on the
`aircraft.
`derived GPS estimated position for the current phase of the
`light mission. It should be recognized that some interpreta-
`tion may be required dependent upon the GPS receiver’s
`ability to simultaneously receive a plurality of satellite
`information signals as is well-understood in the art.
`However, advancements in the art of 12-channel GPS
`receivers have made it no longer necessary to rely on
`interpolation of data as before.
`‘The allowable integrity alarm limit values may change
`depending uponthe phaseof the flight mission. For instance,
`if a pilot is flying in the terminal phase, the integrity alarm
`limit maybeless stringent thanif the pilot is in the approach
`phase of the flight mission. If the pilot is to transition from
`the terminal phasc to the approach phase,the pilot necds to
`know whetherthe current position error boundis sufficient
`to allow the pilot to rely upon the GPS solution to make the
`transition.
`
`10
`
`15
`
`A problem, however, with flight management systems
`employing GPS and IRSis the questionable integrity of the
`GPSestimated position information during those times in
`which RAIM integrity limit values are no longer available,
`ie. insufficient satellite information to provide useful integ-
`rity position error bound values.
`
`BRIEF DESCRIPTION OF THE INVENTION
`
`The present invention uses the FMSto calculate a poisi-
`tion error at anytime, t, after the loss of integrity at time, T,,
`by utilizing the GPS values for position and velocity at time
`T, just before loss of integrity and by utilizing the IRS
`position modified by the known error in IRS position as it
`varies with time and the position error as calculated by the
`GPSvelocity extrapolated over the time (t-T;,).
`BRIEF DESCRIPTION OF THE DRAWINGS
`
`FIG. 1 is a combined inertial reference system and a
`satellite positioning system known in the priorart.
`FIG. 2 is a combined inertial reference system and satel-
`lite positioning system in accordance with the present inven-
`tion.
`
`30
`
`FIG. 3 is a block diagram illustrating a second orderfilter.
`
`DESCRIPTION OF THE PREFERRED
`EMBODIMENT
`
`40
`
`45
`
`50
`
`55
`
`60
`
`65
`
`Illustrated in FIG. 1 is a simplified block diagram of a
`mobile unit,
`a hybrid IRS/GPS estimator commonly
`employed onaircraft. A position estimator 110, for example,
`as part of a flight management system as described earlier,
`receives as inputs (i) GPS output information identified as
`DATA_GPS, further identified by numeral 122, from a
`satellite positioning system receiver 120, and(ii) an inertial
`reference output information indicated by DATA_IRS, fur-
`ther identified by numeral 132 from inertial reference system
`130. Position estimator 110 processes DATA_IRS and
`DATA_GPSto derive (i) a hybrid position estimate iden-
`tified as POS__HYB,and position error estimates identified
`as POS_X_ERR and POS_Y_ERR. This information,
`which will be described in detail below, is provided on
`output signal lines 112, 114, and 116 respectively.
`As is well-understood in the art, satellite positioning
`system receiver 120 includes a satellite signal receiver
`portion 124 for receiving satellite information signals from
`a plurality of satellite vehicles, for example, SV1 and SV2,
`which form,in part, a constellation ofsatellite vehicles. One
`example, already indicated, is the NAVSTAR GPSconstel-
`lation of satellites. In turn, the satellite information signals
`are operated on by GPSposition/solution information pro-
`cessor 126 for providing a GPS solution information iden-
`tified as DATA_GPSonsignal line 122. This information is
`provided as an input to position estimator 110, to a GPS
`integrity validation means 170 and to a GPS RAIM proces-
`sor 150 on commonsignal line 122.
`Inertial reference system 130 includes a plurality of
`inertial sensors indicated by block 131 as inputs to an IRS
`position/inertial information processor 134 for providing
`
`As is well understood in theart, inertial reference systems
`employ a plurality of inertial sensors, for example, gyro-
`scopes and accelerometers, for determining an IRS esti-
`mated position of the aircraft, hereinafter referred to as
`“POS_IRS”. Generally,
`the IRS estimated position is in
`terms of latitude and longitude (altitude being separately
`determined by other means such as an altimeter of some
`type). However, inherent in such inertial sensors are par-
`ticular bias and drift terms which affect the accuracy ofthe
`IRS estimated position of the aircraft utilizing solely an
`inertial reference system. Since high inertial grade sensors,
`Le., low bias and drift characteristics, are very costly, it is
`desirable to minimize the cost of the IRS system byusing
`lower grade inertial sensors.
`In the art, a compromise has been reached by using lower
`grade inertial reference systems in combination with a
`global positioning system to produce a high quality--lower
`cost navigation andflight control system. This is sometimes
`referred to as a Hybrid INS/GPSor IRS/GPSInertial Ref-
`erence System. These systems achieve excellent results
`since low grade inertial reference systems produce very
`accurate dynamic response characteristics, whereas, GPS
`provides very accurate static position information, but less
`accurate dynamic response information. Combining both the
`IRS estimated position and inertial reference information
`with GPSestimated position information provides excellent
`user position information for flight navigation and flight
`control applications. Accordingly, a flight managementsys-
`tem (FMS), combines the excellent features of both the IRS
`and the GPS systems to provide position and inertial refer-
`ence information which permits excellent
`flight
`management, flight control and navigation.
`An example of a hybrid IRS/GPS system is Honeywell
`Inc.’s “Global Positioning Inertial Reference Unit (GPIRU)
`identified as an HG 1050 AGO1 which is referred to as a
`
`“hybrid” system since it provides position and inertial
`information which are a resultant combination of GPS and
`inertial reference system information. The GPIRU includes
`an inertial reference unit with gyros and accelerometers to
`provide information about aircraft attitude and rate of
`change of position as well as providing a first source of
`position information. The GPIRUalso receives inputs from
`
`

`

`5,923,256
`
`5
`IRS derived position and inertial information on signal line
`132, designated DATA_IRS. ‘This information is provided
`as an input to position estimator 110.
`Position estimator 110, which forms in part a flight
`management system known in the art, utilizes the GPS
`solution information, DATAGPS, as a continuousrefer-
`ence for enhancing the accuracy of the IRS position/inertial
`information DATA_IRS, particularly for minimizing result-
`ant bias terms which are inherent in the inertial sensors 131.
`
`Position estimator 110 may also include an input for receiv-
`ing radio position information indicated by numeral 118, e.g.
`VORsignal information designated as DATA.RADIO.
`Position estimator 110 provides as an output on signal line
`112 an estimated position identified as POS_HYB. The
`output POSHYBofposition estimator 110 is provided as
`an input to flight control block 160, useful for deriving
`aircraft flight control signals to achieve a desired aircraft
`position. Vor example, flight control 160 may be employed
`for en-route navigation, terminal approach, and landing of
`an aircraft.
`
`10
`
`15
`
`6
`means 170 receives as its input the GPS receiver output on
`signal line 122 for making such determination, i.e. RAIM
`integrity monitoring available or not available, and provid-
`ing such indication as signal “V” on signal line 172.
`Asis well understood in the art, GPS integrity validation
`means 170 represents a simple analysis of the number of
`satellite information signals tracked by the GPS signal
`receiver 120 which meet predeterminedcriteria, e.g., eleva-
`tion angle. As described earlier, RAIM availability is con-
`ditioned upon having at
`least five satellites tracked for
`receiving satellite information therefrom. Secondly, GPS
`RAIM processor 150 is generally operable not to utilize
`satellite information from those satellites which are less than
`
`a selected elevation angle. In this situation, even though a
`proper numberofsatellites have been tracked, the informa-
`tion may not be reliable due to the elevation angle of the
`satellite relative to the user’s position. In either case, the
`function of GPS integrity validation means 170 is to provide
`an indication of the “non-availability” of RAIM integrity
`limit values, and is provided as an input to a pilot alert
`mechanisms indicated by block 180.
`It should be noted that blocks 150, 155, and 170 are
`shown as discrete functional blocks, for explanation pur-
`poses. However, it should be understood that they may be
`incorporated together, and may also be part of the GPS
`receiver, itself, as should be appreciated by those skilled in
`the art.
`
`it should be noted, as commonly
`Before proceeding,
`understood in the art, that position estimator 110 employs
`filtering techniques, such as second orderfilters or Kalman
`filters for deriving the aforesaid output information. The
`position error estimates POS_X__ERR and POS_Y_ERR
`represent the latitude and longitude errors which are related
`to the differences between the GPS derived position identi-
`
`fied as POS__GPSandthe inertial reference system derived Reliance upon the system as described in FIG. 1 byapilot
`30
`position identified as POS_IRS associated with DATA_
`is extremely dependent upon RAIM availability. In other
`GPS and DATA_IRS, respectively.
`words, the user position estimate POS_ITYB is only useful
`during those times in which RAIMintegrity limit values are
`Further, it should be noted that associated with outputs
`available. Loss of RAIM will have adverse consequences,
`DATA_GPS, DATA_IRS, POS_HYB, POS_X_ERR,
`for cxamplc, requiring the pilot to abort a terminal approach
`and POS__Y__ERRare discrete time values. Accordingly,
`or landing.
`system timing (not shown) and/or interpolation or extrapo-
`lation functions are, of course required, so that position
`For example, consider the situation in which RAIM is
`esimator 110 combines the GPS and IRS information for
`available and an aircraft has already begun the terminal
`phaseofthe flight mission prior to landing. Assume now that
`during this phase of the flight mission that the constellation
`of satellites changes to a condition in which RAIMintegrity
`monitoring is no longer available. In this scenario, the pilot
`is alerted via a warning display mechanization 180 or inputs
`to the flight control system 160 such as to cause the pilot to
`disengage the flight control system which responds prima-
`rily to the aircraft position POS_HYBsince GPS data may
`no longer be reliable. In this scenario, depending upon the
`weather conditions, namely, cloud cover and the like, the
`pilot must determine whether to manually fly the aircraft, or
`abort
`the phase of the flight mission in which RAIM
`integrity monitoring waslost, i.e., not available. In the latter
`case, the pilot may take appropriate actions to require some
`delay time at which the constellation of satellites would be
`in proper position to provide RAIM integrity monitoring
`availability.
`It should be noted, one technique for avoiding the above
`scenario, is predictive RAIM. Predictive RAIM attempts to
`knownin advance that RAIM integrity monitoring is avail-
`able before cntcring a particular phase of the flight and
`would be available throughout the entirety of the phase of
`the flight mission. This is particularly important
`in the
`approach and landing phases of the flight mission. If pre-
`dictive RAIM indicates “non-availability”,
`the pilot may
`take certain actions, e.g., decrease the aircraft velocity such
`that landing takes place at a later time when RAIM is once
`again available.
`Illustrated in FIG. 2 is one embodiment of the present
`invention for providing an alternate GPS integrity limit
`
`In the following
`substantially the same time values.
`exposition, synchronization of time values should be
`assumed and that each value has a discrete time associated
`therewith.
`
`As is well-understood in the art, the GPS position solution
`information must be validated by a GPS system integrity
`monitor. GPS RAIM processor 150is intended to operate on
`the GPSsolution information DATA_GPSfor determining
`at least horizontal integrity limit values HIL, and may also
`provide vertical integrity limit values VIL. In turn, these
`integrity limit values are compared in RAIM Integrity Limit
`Comparator 155 with selected integrily alarm limit values
`dependent upon the phase of the flight mission. In turn, if
`HIL/VILis acceptable, the pilot will allow control of the
`aircraft based uponthe outputs of position estimator 110. On
`the other hand, if HIL/VIL exceeds the integrity alarm limit
`values, the pilot must be alerted so that corrective action
`may be taken.
`Asecondscenario is, of course, the case where RAIM is
`unavailable, i.e., insufficient number oftracked satellites. In
`this scenario, the constellation of satellites as observed by
`the GPS receiver 120 may be such thatit is impossible for
`GPS RAIMprocessor 150 to arrive at a solution for obtain-
`ing HIL and/or VIL integrity limit values—these being
`resultant large values for HIL/VIL. Accordingly, GPS integ-
`rity validation means 170 is employed to provide an indi-
`cation as to whether or not there exists RAIM integrity
`monitoring availability, i.c., sufficient satellite information
`signals to be able to calculate the integrity limit values, HIL
`and/or VIL.As illustrated in FIG. 1, GPS integrity validation
`
`35
`
`40
`
`45
`
`50
`
`55
`
`60
`
`65
`
`

`

`5,923,256
`
`7
`process during those time periods in which the usual RAIM
`integrity monitoring is not available. In FIG. 2, similar
`functioning blocks as those illustrated in FIG. 1 have
`retained the same numeral designation, and therefore will
`not be further described. FIG. 2, in addition to those com-
`ponents as shown in FIG. 1, further includes velocity/
`acccleration crror estimator 210, modificd IRS position
`estimator 220, position difference calculator 230, alternate
`integrity limit comparator 240, and position selector 300.
`
`10
`
`it should be understood that GPS
`Before proceeding,
`position/inertial
`information processors may provide
`position/inertial information in a variety of coordinate ref-
`erence frames. Commonly, GPS information provides posi-
`tion information in an earth-centered, earth-fixed, coordinate
`reference frame. In turn, this information may be translated 1
`into latitude and longitude values. The inertial GPS solution
`information may include velocity information in terms of
`north direction and east direction, as is commonly found in
`the art. These values, of course, are mere translations and/or
`transformations of the earth-centered, earth-fixed, position
`information. Accordingly, as depicted in FIG. 2 and the
`explanation which follows, north and east directions are
`represented by X and Y, respectively, which also relates ta
`latitude and longitude, respectively. P'urther, in the following
`exposition, the term “POS”represents position, and the term
`“VEL” represents velocity.
`
`Referring again to FIG. 2, velocity/acceleration error
`estimator 210 receives as inputs (i) GPS derived velocity
`information in the X and Y direction, and (ii) IRS velocity
`information in the X and Y direction derived from the
`inertial sensors, are provided as inputs. These terms may be
`represented by:
`
`VEFI._X_GPS
`VEL__Y__GPS
`VEL__X_IRS
`VEL__Y_GPS
`
`GPS derived velocity, X direction
`GPS derived velocity, Y direction
`IRS derived velocity, X direction
`IRS derived velocity, Y direction
`
`where the IRS and GPS designate refers to data derived from
`the inertial reference system 130 and the GPSreceiver 120
`respectively. As before, these terms have substantially iden-
`tical corresponding real time values associated therewith.
`Velocity/acceleration error estimator 210 provides as an
`output information designated as “DATA_ERROR”which
`represents discrete acceleration and velocity error, or bias
`terms in the IRS position/inertial information. Such terms
`may be represented by VEL_X_ERR, VEL_Y_ERR,
`ACC_X_ERR, and ACC__Y__ERR, X and Y velocity and
`acceleration errors, respectively. Modified IRS position esti-
`mator 220 receives as inputs DATA_ERROR,the inertial
`reference system 130 position information represented by
`POS_IRS, the position errors POS_X_ERR and POS_
`Y_ERR,andvalidation signal on signal lines 332, 134, 114,
`116, and 172, respectively.
`Modified IRS position estimator 220 is intended to pro-
`vide an output on signal line 222 representative of a modi-
`fied IRS position estimate designated as POS_IRS_LOSS
`which represents an estimate of the real position of the user
`during the time period in which GPS RAIM integrity
`monitoring was not available (1.c., RAIM “LOSS”) follow-
`ing a time period when RAIMintegrity monitoring was
`available. Modified IRS position estimator 220 is intended
`to operate on the aforesaid input information for determining
`a position estimate which maybe mathematically described
`as follows:
`
`30
`
`35
`
`40
`
`45
`
`50
`
`55
`
`60
`
`65
`
`8
`
`POS_IRS_LOSS(t) = [POS_IRS()] —
`
`[POS_ERR|to | + [VEL_ERR(to)|[t - T,] + [.5][ACC_ERR(to)})|[r- nl
`
`These terms, of course, having their coordinate
`components, i.e. X and Y. The above expression is simply a
`statement that the modified IRS estimated position POS__
`IRS_LOSSat time “t” is the measured IRS position POS__
`IRS(t) corrected by the velocity and acceleration error terms
`provided by Velocity/Acceleration error estimator 300 and
`the position error estimates provided as an output of position
`estimator 110—the latter being provided bya filtered error
`estimator described earlier.
`
`In component terms, then
`
`D= Rex (PX PX)? + (PY — PX)? x COSUATV)
`
`x,y = latitude, longitude,position coordinates
`
`where
`
`Pi = POS_X_GPS
`Pi = POS_Y_GPS
`Py = POS_Y_IRS
`PY = POS_Y_IRS
`LATV = Ath2
`RE = Earth Radiusat LATAV
`
`the modified IRS position
`Again referring to FIG. 2,
`estimate identified as POS_IRS__LOSSprovided on output
`signal line 222 is presented to position difference calculator
`230. Position difference calculator 230 receives as a second
`
`the GPS derived position identified as POSGPS
`input
`provided on output signal line 129 from GPSreceiver 120.
`Position difference calculator 230 is intended to derive the
`difference in position D between the GPS derived position
`and the modified IRS position estimate, as mathematically
`described by:
`
`D= Re #y(Pj — PS? + (P} — P} «COSUATV)
`
`x,y = latitude, longitude, position coordinates
`
`where
`
`P| = POS_X_GPS
`Pi = POS_Y_GPS
`P, = POS_Y_IRS
`Py = POS_¥_IRS
`
`LATV = Pit Ps
`2
`RE = Earth Radiusat LATAV
`
`It should be noted that the position difference “D” repre-
`sents the magnitude of a vector between (a) the position
`coordinates derived from the GPSsolution provided by GPS
`receiver 120, and (b) the position coordinates derived by the
`IRS position/inertial information processor 134 as modified
`by the velocity/acceleration errors DATA ERROR pro-
`vided as an output of modified IRS position estimator 220,
`namely POS_IRS_LOSS. The value “D” represents an
`“alternate integrity limit” value since it relates to the errors
`in the IRS system 130 at the time RAIM integrity monitoring
`waslost. The value D is provided as an output on signalline
`
`

`

`5,923,256
`
`9
`232, and presented as an input to alternate integrity limit
`comparator 240.
`Alternate integrity limit comparator 240 receives as inputs
`the alternate integrity limit value D, and integrity alarm limit
`reference values identified as reference-I and reference-I],
`dependent upon the phase of the flight mission. Alternate
`integrity limit comparator 240 is intended to compare the
`deviation between the alternate integrity value D and a
`predetermined flight phase integrity alarm limit value(i.e.,
`the alarm limit reference values). The aforesaid integrity
`alarm limit value is of course dependent upon the phase of
`the flight mission (for example, terminal phase, approach
`phase,or final approach (landing) phase). Alternate integrity
`limit comparator 240 provides an indicat