DIFFERENTIAL SATELLITE POSITIONING SYSTEM GROUND STATION WITH INTEGRITY MONITORING

Abstract: In the present invention, a plurality of DGPS receivers are employed to determine receiver-satellite-specific differential correction values. In turn, satellite-specific differential correction values are determined as an average of the individual receiver-satellite-specific differential correction values for each of the tracked satellites. A discriminator is employed for determining a receiver-satellite-specific discriminant which may be sampled and subsequent determination of a receiver-satellite-specific statistical value thereof. The receiver-satellite-specific statistical values may then be employed for determination of a Detection Threshold value and other integrity information associated with each of the satellite-specific differential correction values. The ground station further includes a means for communicating the satellite-specific differential correction values and integrity information associated therewith to a DGPS airborne receiver. A DGPS airborne system in accordance with the present invention is configured to include a mechanization for receiving the satellite-specific differential correction values and integrity information associated therewith. An airborne integrity monitoring system may then be employed to derive particular integrity limit values associated with the DGPS derived position for pilot awareness as to whether or not the pilot should rely upon the DGPS solution for a precision landing.

Description: DIFFERENTIAL SATELLITE POSITIONING SYSTEM GROUND STATION WITH INTEGRITY MONITORING BACKGROUND OF THE INVENTION Field ofthe Invention The present invention is related to differential GPS systems, and more particularly, to a DGPS system employing integrity monitoring. Description ofthe Related Art Satellite positioning systems, commonly referred to as global positioning systems or simply GPS, are now well-known in the art. Such specific systems, for example, NAVSTAR-GPS, are rapidly being used for determination of the geocentric position of mobile units, such as water and land vehicles, aircraft and survey equipment, to name a few. In aircraft, GPS systems are being utilized for navigation, flight control, and air space control. These GPS systems may operate independently, or in combination with, among others, inertial reference systems or attitude heading reference systems in order to provide information during an aircraft flight mission. Global positioning systems similar to NAVSTAR commonly use a GPS receiver, located on a mobile unit, for receiving satellite information signals transmitted from a plurality of satellites. Each GPS satellite transmits an 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 distances and knowledge ofthe satellite positions are then used to compute the position ofthe receiver unit using known triangulation techniques. For example, in the NAVSTAR-GPS system, a mobile unit with a GPS receiver, such as an aircraft, detects a pseudo random code contained in a given GPS satellite information signal and derives therefrom the "elapsed time" or time delay between the transmission ofthe satellite signal and its reception at the GPS receiver. From this time delay, the GPS receiver derives the range between the GPS receiver antenna and the satellite, sometimes referred to as the pseudo range or pseudo range measurement. Herein, the GPS receiver's position, or the mobile unit's position, generally refers to the corresponding antenna position. In addition, as part ofthe NAVSTAR-GPS system, each satellite information signal also contains precise ephemeris data and course almanac data which both describe the corresponding satellite orbital trajectory in earth centered space as is well known in the art. The coordinates ofthe satellite's orbital position at the transmission time may be derived from either the ephemeris data or the course almanac data. The geocentric position ofthe satellite may be calculated with a higher degree of precision from the ephemeris data than is possible with the almanac data as is well-known It should be understood that the mobile unit ' s three-dimensional geocentric position coordinates are referenced to the World Geodetic Coordinate System. Herein, it should be recognized by those skilled in the art that the World Geodetic System is an earth-centered, earth-fixed geocentric coordinate system, which may be converted to any other coordinate system as required by the user. Sometimes the aforementioned coordinate system is referred to as the WGS84 earth-centered, earth-fixed, rectangular coordinate frame. Herein, the World Geodetic System Coordinates should be presumed, and position refers to this three dimensional WGS84 coordinate system. In order to determine the position ofthe GPS receiver unit, a minimum of four satellite signals are required, rather than the expected three. This is so, since the GPS receiver includes a receiver clock which is not as accurate as the atomic clock ofthe satellites. Therefore, receiving satellite infoπnation 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. Herein, the corrected receiver clock time is referred to as the receiver time. Thus, if signals from four or more satellites are available to the GPS receiver unit, the geocentric position ofthe receiver may be determined within approximately one-hundred meters of its "true" geocentric position. Herein, the receiver position derived by the triangulation technique using data from multiple satellites is referred to as the "estimated position". The accuracy ofthe estimated position ofthe receiver unit is dependent upon many factors including, among others, atmospheric conditions, selective availability, and the line of sight view ofthe satellites. Although the satellite positioning system referred to as GPS is by far the most accurate global navigation system ever devised, its accuracy can be boosted using a technique called "differential GPS", sometimes referred to as "DGPS". DGPS can achieve measurement accuracy better than a meter. Differential GPS has been widely used in surveying applications, and now its use is being developed for aircraft approach and landing applications. With regard to the latter, GPS systems standards have been developed by the RTCA Inc. (formerally called Radio Technical Commission for Aeronautics) in association with aeronautical organizations ofthe United States from both government and industry. The RTCA has defined performance requirements for a DGPS system as particularly identified in Document No. RTCA/DO-217, dated August 27,1993; and for GPS performance requirements for navigation equipment as particularly identified in Document No. RTCA/DO-208, dated August 27,1993, both of which are incorporated herein by reference thereto. As is well understood in the art, a differential GPS system incoφorates a reference or "ground station" which includes a GPS receiver's antenna installed at a surveyed site. The geocentric position ofthe GPS receiver is known from surveying techniques. The GPS ground station receiver determines pseudo range values between the receiver and a plurality of satellites. Since the position ofthe satellite is derived from the satellite data associated with the satellite signals, and the position ofthe receiver is also known, a calculated range value therebetween may be determined for each ofthe tracked satellites. In turn, the difference between the measured pseudo range value and the calculated range value for each of the tracked satellites may be determined. This difference is commonly referred to as the "differential correction". The differential correction value is essentially the pseudo range error between the "observed" or "measured" pseudo range value derived from the satellite signal travel time and the calculated range value between the antenna's position and corresponding satellite position. The motivation for differential operation is that many ofthe largest GPS error sources, such as selective availability and ionospheric effects, are common to two or more receivers operating in spatial and temporal proximity, since these anomalies affect the satellite signal travel time. These error sources can be nearly eliminated in the differential mode, by determining and applying the differential correction value or pseudo range error for greatly enhancing performance. A differential GPS system is shown and described in a publication entitled, "Design and Flight Test of a Differential GPS/Inertial Navigation System for Approach/Landing Guidance", Navigation: Journal of Institute of Navigation, Summer 1991, Vol. 38, No. 2, pp 103-122, incorporated herein by reference thereto. As described therein, these differential corrections may be transmitted, by any data link technique, to a mobile unit which incorporates a GPS receiver and associated antenna experiencing substantially the same errors in the pseudo range observed values for the same tracked satellites as the ground station receiver. Accordingly, the mobile unit's pseudo range observed values may be corrected by the differential corrections for a more precise determination ofthe geocentric position ofthe mobile unit. It should be noted that the accuracy ofthe corrections is somewhat dependent on the proximity of the mobile unit relative to the ground station. Although differential GPS ground stations have advanced the art of GPS for some applications, there is a need for a differential GPS ground station with enhanced integrity, continuity, and availability as particularly identified in the aforesaid RTCA publication. SUMMARY OF THE TNVENTJON An object ofthe present invention is to provide integrity information associated with a DGPS ground station which may be incoφorated by an aircraft for a precision landing system. In the present invention, a plurality of DGPS receivers are employed to determine receiver-satellite-specific differential correction values. In turn, satellite- specific differential correction values are determined as an average ofthe individual receiver-satellite-specific differential correction values for each ofthe tracked satellites. A discriminator is employed for determining a receiver-satellite-specific discriminant value which may be sampled to allow subsequent determination of a receiver-satellite- specific statistical value thereof. A discriminator value, computed as the differences between eacah receiver-satellite-specific differential correction and the average satellite- specific differential correction, is compared to a detection threshold. The detection threshold is set based on statistical values that have been determined using previously collected receiver-satellite-specific discriminant values. The receiver-satellite-specific statistical are also utilized for obtaining other integrity information associated with each ofthe satellite-specific differential correction values. The ground station further includes a means for communicating the satellite-specific differential coπection values and integrity information associated therewith to a DGPS airbome receiver. A DGPS airbome system in accordance with the present invention is configured to include a mechanization for receiving the satellite-specific differential coπection values and integrity information associated therewith. An airbome integrity monitoring system may then be employed to derive particular integrity limit values associated with the DGPS derived position for pilot awareness as to whether or not the pilot should rely upon the DGPS solution for a precision landing. BRIEF DESCRIPTION OF THE DRA WINGS Figures 1 A- IB is a block diagram illustrating a prior art GPS signal receiver. Figure 2 is a block diagram illustrating a prior art ground station differential GPS system and an airbome differential GPS system. Figure 3 is a schematic diagram illustrating a DGPS ground station system in accordance with the present invention in conjunction with an aiφort runway. Figure 4 is a block diagram illustrating the differential coπection processing function of Figure 3. Figure 5 is a functional block illustrating a ground station integrity monitoring system in accordance with the present invention. Figure 6 is a block diagram illustrating an airbome DGPS signal receiver system in accordance with the present invention. DESCRIPTION OF THE PREFERRED EMBODIMENT Illustrated in Figure 1 is a prior art GPS signal receiver 100 having an associated receiving antenna 102. Herein, the acronym "GPS" generally refers to any satellite- based positioning system wherein each of a plurality of satellite vehicles employs a transmitter for transmitting a satellite information signal from which a user may determine the user's position derived from a satellite information signal transmitted from at least four satellite vehicles. Such satellite-based positioning systems include at least those commonly refeπed to as "GPS" and "GLONASS". The prior art GPS signal receiver 100, as well understood in the art of satellite positioning systems, is always coupled to a GPS receiving antenna, and serves to determine the actual position ofthe receiving antenna. Herein, the position ofthe antenna will be assumed even though phraseology may incoφorate the use ofthe "position ofthe receiver" or "user position". As illustrated in Figure 1, the primary components of GPS signal receiver 100 are satellite signal receiving means 110 for receiving the transmitted satellite information signals, and satellite signal tracking and information processing means 120 which serves to decode all ofthe information contained in the satellite information signal in accordance with those standards specified by a particular satellite positioning system, for example, those specified by the U.S. Department of Defense with regard to the satellite positioning system commonly refeπed to as "GPS". More particularly, satellite signal tracking and information processing means 120 generally includes hardware/software for decoding (i) satellite-specific coded information associated with each transmitted satellite information signal received by signal receiving means 110, (ii) selected coded information (gold code) from which each satellite infoπnation signal transit time may be derived, and (iii) selected carrier signal parameters associated with each satellite information signal (e.g., frequency, phase, etc.) as received by satellite information signal receiving means 110. The aforesaid information is depicted in Figure 1 by receiver-satellite-specific information block 130, receiver-satellite-specific measurement information block 140, receiver-satellite-specific carrier signal measurement information block 150, respectively. Associated with block 130 is the common receiver-satellite-specific information, including, but not limited to Satellite Trajectory Data (commonly, the emphemeris data), and: SVID = satellite vehicle identification number, IOD = issue of data (satellite clock time at time of transmission of the information signal), and Toe = time of applicability. Receiver-satellite-specific measurement information block 140 represents generally, the information derived from coded information coded onto the caπier signal ofthe satellite information signal, namely: tm = receiver-specific measurement time, P n = pseudo range value, and S/No = receiver signal to noise ratio R at (t m ) = satellite position at the receiver measurement time calculated from the satellite trajectory information. Receiver-satellite-specific caπier signal measurement information block 150 is representative ofthe information derived from the carrier signal characteristics of each ofthe satellite information signals, and includes, among others: f = satellite information signal frequency as observed by the receiver, φ = phase measurements at the receiver, ΔP = delta range value, Doppler-Shift = change in frequency of the caπier signal as observed by the receiver between two close points in time, and Integrated caπier count = count information which represents the phase change ofthe received satellite-specific information signal observed by the receiver between two close points in time. Before proceeding, it should be understood that each measurement derived from the received or tracked satellite information signal is valid for a specific receiver measurement time. For multi-channel receivers, e.g. 12 channels, the GPS signal receiver may provide concuπent measurements, specifically receiver-satellite-specific pseudo range values and delta range values, and other information as aforesaid. However, for this data to be useful, all of this data must be related to the common receiver-specific measurement time " t m ". Different GPS signal receivers, even though they may be tracking the same satellite vehicles, may have a different receiver-specific measurement time primarily because of different internal clock times, noise factors, and possible multi-path effects. In the following exposition, and particularly the mathematical expressions which follow, it should be assumed that information derived from a particular satellite information signal by a particular receiver will have associated therewith an implied receiver-satellite-specific measurement time t m even though not expressed. The GPS signal receiver 100 as briefly described above is commonly available and provided by such manufacturers as Honeywell Inc., Trimble, Novaltel, and Collins Radio, among others. GPS signal receiver 100 as described is generally embodied by sophisticated analog circuits and subsystems, digital networks, and computational devices including, among others, microprocessors, computers, and the like. As is well understood in the art, computational devices specifically embodied in the GPS signal receiver 100 may also be part of other subsystems for providing multiple or multiplexed tasks for other operations as is well known. Further illustrated in Figure 1 is a block diagram illustrating further functions commonly associated with the GPS signal receivers ofthe prior art. First, thereshown is caπier smoothing block 160 responsive to the pseudo range value P n and the delta range value ΔP , thereby providing a "caπier smoothed" pseudo range value P n as described in a publication entitled, "GPS Doppler Processing for Precision Positioning in the Dynamic Application, by Jarad Ashjaee, Nov. 1985, IEE Oceans Conference. It should be noted, however, that the function of caπier smoothing may be performed in other subsystems as described. GPS signal receiver 100 may also include user-position processing identified by block 170 responsive to pseudo range values, or the caπier smoothed pseudo range values as illustrated in Figure 1, from a plurality of satellites for determining the user's position U(x, y,z) by triangulation as is well understood in the art. In turn, position solution matrix processing may be incoφorated as identified by block 175 to provide a position solution (least square) matrix S£CEE[4 x N] which is used to obtain the user's position relative to all ofthe acquired satellite information signals, where the four dimensions represent the three axes ofthe coordinate reference frame and time, as is well-known. The GPS signal receiver 100 may also include further refinements including an azimuth/elevation block 180 and tracking monitor 190. Azimuth/elevation block 180 may serve to first determine the azimuth/elevation angle of each ofthe tracked satellites relative to the user's position for subsequent determination as to whether or not the satellite information signals should be used, or not, based on minimum elevation angle selection criteria, commonly, 5°. Tracking monitor 190 may be included to provide an aπay of information indicating which ofthe satellites, by satellite vehicle identification number, "n", is considered to be an "acquired" satellite, or a "lost" satellite based on the azimuth/elevation angle selection criteria, signal reception, if any, or other criteria, as imposed by the user, generally embodied in software/firmware. Again, referring to Figure 1, GPS signal receiver 100 provides, as outputs, the receiver measurement time t m , the satellite vehicle identification number, hereinafter refeπed to as simply "n", the satellite position Rc τ(^w) > *h e pseudo range value p n , which may be carrier smoothed as aforesaid, delta range ΔP , as well as the GPS signal receiver (user) position U(x,y,z) , and coπesponding position solution matrix S^CEE ■ These measurements may be further employed in a differential GPS system as will briefly described herein, and also described in a publication entitled Design and Flight Test of a differential GPS/Inertial Navigation System for Approach/Landing Guidance, Vallot et al., Joumal ofthe Institute of Navigation, Vol. 38, No. 2, Summer 1991. The GPS signal receiver 100 briefly described above may be employed in a differential GPS (DGPS) aircraft landing system as particularly illustrated in Figure 2. The system illustrated in Figure 2 is similar to that as generally described in U.S. patent 4,894,655, entitled "Landing Assistance System Using Navigation Satellites", issued to Jognet et al., and U.S. patent 5,361,212, entitled "Differential GPS Landing System", issued to Class et al. Refeπing to Figure 2, thereshown is a DGPS ground station 200 and a DGPS airbome system 250, the latter being intended to be part of an aircraft flight control and/or navigation guidance system. Both DGPS ground station 200 and DGPS airbome system 250 each include a similar functioning GPS signal receiver 100, like that of Figure 1 , and are such that when the DGPS airbome system 250 is in proximity to the DGPS ground station 200, for example within 50 miles ofthe ground station, their respective GPS signal receivers 100 will generally track a common set of satellite vehicles, for example , SV 1 , SV2 , ...SV n . The superscript "n", herein, refers to specific tracked ones ofthe constellation of "N" satellite vehicles of a particular satellite positioning system. DGPS ground station 200 further includes differential coπection processing designated by block 210, an up-link transmitter 220, and transmitting antenna 222. Differential coπection processing block 210 generally includes a calculated range processing block 212 for determining receiver-satellite-specific calculated range values X n as a function ofthe difference between the known surveyed position of GPS signal receiver 100 associated with ground station 200, represented by G (x,y,z), and each of the tracked satellite positions, provided by GPS signal receiver 100 and represented by ^SAT as s we ^ understood. In turn, differential coπection calculation block 214 is operative for determining a satellite-specific differential coπection value C L , as a function ofthe satellite specific pseudo range value P n ' and the calculated range value X n , and the estimated receiver-specific offset or clock bias (not shown) as is well known. These satellite specific coπection values C are transmitted to the DGPS airbome system through up-link transmitter 220 and antenna 222. As is well understood in the art, all GPS signal receiver measurement values have a receiver-satellite-specific measurement time, t m associated therewith. Further, it should be noted that the subscript "wob" implies that the GPS signal receiver clock bias has been taken into consideration. That is, the differential coπection value has been adjusted for any eπor caused by clock bias. Hereafter, the satellite-specific differential coπection values C will be represented by C n . Again referring to Figure 2, DGPS airbome system 250 is illustrated as generally including a DGPS signal receiver 252 and a flight guidance system 284 which includes lateral/vertical deviation processing block 285 and glide path information block 288. DGPS signal receiver 252 includes a GPS signal receiver 100, like that of ground station 200. DGPS signal receiver 252 further includes pseudo range coπection processing indicated by block 260. Airborne system 250 further includes differential coπection infoπnation receiver 290 operative for receiving the satellite specific coπection values C n , and in turn provide them as information to pseudo range coπection processing block 260 as is indicated by data signal line 296. DGPS signal receiver 252 is operative for providing the airborne user position A(x,y, z) and the position solution matrix S£CEE[4 x N _ indicated by outputs 254 and 256, respectively. This information is based on the measurement information provided by the airbome GPS signal receiver 100 coπected by the differential coπection data C n , and may be generally derived by widely known Kalman filtering techniques as well as least square type algorithms, and the like. Airbome flight guidance system 284 is intended to be responsive to the DGPS derived aircraft position A(x, y,z) and the position solution matrix S£CEE[4 x N] provided as outputs of DGPS signal receiver 252 for providing an indication of any deviation between the aircraft position and an intended position, for example along a runway precision approach glide path. This is illustrated in Figure 2 by lateral/vertical deviation processing block 285 intended to provide the appropriate lateral/vertical deviation eπor signals to an auto-pilot 289 so as to cause the aircraft to descend along a prescribed glide path provided by glide path infoπnation block 288. The glide path information 288 may be either stored in memory, or transmitted to the aircraft from the ground station as taught in the aforesaid patents, and also depicted in Figure 3 by block 390 as will be further described. Illustrated in Figure 3 is a schematic diagram illustrating a DGPS ground station in accordance with the present invention and intended for employment with an aircraft for a precision approach landing system at an aiφort. Thereshown are a pair of mnways 305 and 310. In close proximity to the runways are three remote GPS signal receivers 100 identified as receivers 100A, 100B, and lOOC, substantially similar to GPS signal receiver 100 of Figure 1. GPS signal receivers 100 A-C are intended to be spatially located such that their respective antennas are in spaced apart aπangement, being approximately 100 meters from each other, so as to separately receive GPS satellite information signals at substantially the same time. The use of a plurality of GPS signal receivers for a ground station, in accordance with the present invention, is intended to insure that a strong multi-path signal affects only one ofthe GPS signal receivers. Secondly, if the plurality of GPS signal receivers for the ground station are properly spaced apart, the problem ofthe aircraft, itself, masking GPS satellite information signals from more than one ofthe plurality GPS signal receivers during the landing phase ofthe flight mission may be avoided, thereby diminishing the likelihood of a need to abort the landing due to loss of GPS acquired satellite information signals caused by a loss of a ground station GPS signal receiver redundancy check. Further shown in Figure 3 is a differential coπection processing means 300, in accordance with the present invention, responsive to receiver-satellite-specific information from each ofthe remote GPS signal receivers 100A-C, for deriving satellite-specific differential coπection information and coπesponding integrity information, designated by numerals 310 and 312, which may be transmitted to an aircraft or mobile unit by up-link transmitter 220 and antenna 222, similar to that as shown in Figure 2. Further details ofthe differential coπection processing means 300 will be subsequently described with respect to Figure 4. Before proceeding, it is of paramount importance to understand that each ofthe GPS signal receivers 100A, 100B, and 100C, and coπesponding antennae 102A, 102B, and 102C, respectively, operates totally independent for deriving "receiver-satellite- specific" information associated with as many tracked satellite vehicles as possible. In the exposition which follows, the term "receiver-satellite-specific" will be used to distinguish that the information is associated with only one ofthe GPS signal receiver and derived from one ofthe satellite information signals associated with a specific satellite vehicle. Similarly, the term "satellite-specific" will be used to identify information associated with a specific satellite. Accordingly, in the mathematical expressions which follow, it will be assumed that there are "N" satellites, and "K" receivers. Further, the mathematical terms may have a superscript pair, which designates a specific satellite "n", and specific receiver "k", (k= A,B,C, etc. or k=l ,2,3) so as to obtain the desired receiver-satellite-specific information, or the satellite-specific information. Illustrated in Figure 4 is a block diagram illustrating the differential coπection processing means 300 of Figure 3 in accordance with the present invention. Thereshown is differential coπection processing means 210A receiving inputs G , P ' and R T fr° m G signal receiver 100A, similar to differential coπection processing means 210 and GPS signal receiver 100, respectively, as already described with reference to Figure 2. Differential coπection processing means 210A provides a raw differential coπection output value indicated as C n ' . It should be noted here that the measurement time, " t m " (here, t m ), is always associated with a specific GPS signal receiver, 100 A, 100B, and 100C, and G is the known geocentric position ofthe coπesponding GPS signal receiver 100 A. Further, differential coπection processing means 300 includes differential coπection processing means 210B and 210C, receiving inputs from GPS signal receivers 100B and 100C, respectively, also being substantially the same as differential coπection processing means 210 and GPS signal receiver 100, respectively, as already described, and provides raw differential coπection value outputs designated C n ' and C n ' , respectively. Again, it should be noted that GPS signal receivers 100 A-C each independently derives measurement information as described earlier with reference to GPS signal receiver 100 illustrated in Figure 1. In particular, each GPS signal receiver 100 A-C may inherently have a clock time offset or bias with respect to the atomic clock time ofthe satellite vehicles. Secondly, since the GPS signal receivers are not necessarily tightly synchronized, the receiver-satellite-specific tracking information and the receiver- satellite-specific measurements may have differing measurement times " f m ". Accordingly, differential coπection processing means 300 includes clock bias coπection processing means 410 and, optionally, measurement time synchronization processing means 420. Clock bias coπection processing means 410 provides discrete differential coπection values C ' , which represents the raw differential coπection value "' without clock bias. In turn, these output differential coπection values are presented to measurement time synchronization processing means 420. As indicated earlier, caπier smoothing may be altemately embodied in subsystems other than the GPS signal receivers 100. In the present invention, caπier smoothing is preferably embodied in block 300 such that the representation p n ' in Figure 4 represents raw pseudo range and delta range measurements; and differential coπection processing blocks 210A-C may incoφorate the caπier smoothing function similar to that already described with reference to Figure 1. Measurement time synchronization processing means 420 is responsive to the output ofthe clock bias coπection processing means 410, namely the receiver-satellite- specific coπection values with clock bias removed, namely C n ' , so as to adjust these values to a common synchronization time t S y nc . As illustrated in Figure 4, the output of measurement time synchronization processing means 420 is designated C n ' representing the receiver-satellite-specific differential coπection values with clock bias removed and adjusted to reflect a common measurement time t sync . One example of measurement time synchronization processing means 420 is simply determimng the median or average measurement time associated with each of γt ir the raw differential coπection values C ' , , and adjusting the raw differential coπection values C "' , , to the median or average time, and providing output differential coπection values C ' . This is desirable since associated with each value of C F is potentially a significant differing receiver-satellite-specific measurement times, t^ since the plurality of GPS signal receivers 100A-C are not necessarily tightly synchronized. On the other hand, if all GPS signal receivers are able to make measurements at the same measurement time, then the synchronization processing means 420 would no longer be necessary. Similar to the system of Figure 2, the puφose of differential coπection processing means 300 is to provide satellite-specific differential coπection infoπnation associated with all ofthe tracked satellites meeting the selected criteria as aforesaid, and in which the satellite-specific differential coπection information is a function of receiver-satellite-specific information obtained by a plurality of GPS signal receivers. Accordingly, differential coπection processing 300, in accordance with the present invention, further includes a satellite-specific differential coπection averaging processing means 430 for processing the receiver-satellite-specific coπection values "' so as to derive satellite specific differential coπection average values represented by C n , i.e., the average ofthe satellite-specific differential coπection values observed by each ofthe three GPS signal receivers 100 A-C. This may be accomplished by a simple average or a selected weighted average, or the like. Before proceeding, it should be recognized by the reader that the configuration of differential coπection processing means 300 illustrated in Figure 4 may altematively be arranged within the scope ofthe present invention. For example, each of GPS signal receivers 100 A-C may include the differential coπection processing function and clock bias removal function indicated by designated blocks 210A-C and block 420. That is, each ofthe GPS signal receivers may contain the necessary functions for obtaining the receiver-satellite-specific differential coπection values C n ' > , before any time synchronization, if necessary, and averaging performed by blocks 420 and 430, respectively, or other general functions as will be further described. In accordance with the present invention, it is of paramount importance for a DGPS landing system, when employing DGPS as a primary control input for aircraft flight control, that the integrity ofthe DGPS derived position ofthe aircraft be within prescribed limits, more specifically, as described by the FAA, CAT-I requirements. Reference is hereby made to the following RTCA publications which describe requirements for GPS/DGPS systems: (i) Minimum Operational Performance Standards for Airbome Supplemental Navigation Equipment Using Global Positioning System (GPS), RTCA/DO-208 July 1991, Prepared by: Special Committee - 159, and (ii) Minimum Aviation System Performance Standards DGNSS Instrument Approach System; Special Category I (SCAT-I), RTCA/DO-217 August 27, 1993, Prepared by: Special Committee - 159. Further GPS system requirements are set forth in a FAA document identified as: IFR Approval for Differential Global Positioning System (DGPS), Special Category I, and a technical Order, Instrument Approaches Using Private Ground Facilities Identified as Order No. 8400.1 1 , Department of Transportation, Federal Aviation Administration. September 15, 1994. These aforesaid documents are incoφorated herein by reference. Reference is also made to the The Institute of Navigation Proceedings of ION GPS-90, September 1990, and more particularly to a paper entitled, "Implementation of a RAIM Monitor in a GPS Receiver and an Integrated GPS/IRS, by Mats Brenner. The latter mentioned paper describes in detail a foundation of well accepted techniques for establishing integrity numbers related to aircraft position based on a GPS derived position solution. Portions of this paper forms a foundation ofthe prior art with regard to the term "integrity" as used herein. In order to provide a precision DGPS landing system having the required integrity, the present invention incoφorates ground station integrity monitoring providing two distinct integrity monitoring functions. The first is a discrete data integrity monitoring system intended for deselecting particular ground station receiver- satellite-specific information which is unacceptable, and would contaminate the resulting satellite-specific differential coπection infoπnation C n obtained by the ground station's differential coπection processing means 300. The second is ground station integrity information or peformance processing which establishes satellite- specific statistical integrity parameters related to the results associated with the satellite- specific differential coπection infoπnation, which may in turn, be transmitted to a DGPS airbome landing system and utilized in combination with an airbome integrity monitoring system for subsequent determination ofthe integrity ofthe overall flight control and/or guidance DGPS derived drive signals. The airbome integrity monitoring system in the prefeπed embodiment ofthe invention is intended to establish lateral and vertical integrity limit values which may be compared to selected "maximum" limit values. Based on these comparisons, a pilot or automatic control device may decide as to whether or not to rely upon a guidance signal derived by the DGPS position solution for controlling aircraft precision landing, and which ultimately may require a decision to abort the DGPS landing or use an altemative landing system or manual technique. Illustrated in Figure 3 is a ground station integrity monitoring system 500 associated with the differential coπection processing means for providing integrity output information indicated by numeral 312 as an input to transmitter 220 intended for subsequent transmission to an aircraft. Details of ground station integrity monitoring system are particularly illustrated with reference to Figure 5. As should be understood by those skilled in the art, the ground station integrity monitoring functions may be incoφorated as a component part of differential coπection processing means 300, however are described herein separately for ease of understanding the present invention. Illustrated in Figure 5 is a block diagram illustrating a ground station integrity monitoring system 500 in accordance with the present invention. Ground station integrity monitoring system 500 includes discriminator 510 which serves to provide receiver-satellite-specific discriminant values Z n " , and statistical value processing depicted by block 520 which serves to provide receiver-satellite-specific statistical discriminant values <X . Integrity monitoring system 500 further includes discrete data integrity monitoring system 540 which serves to provide a detection threshold value, DT, and a quantification of which receiver-satellite-specific discriminant value Z n ' exceeds the detection threshold value, DT, namely Fail [n,k]. It should be noted that DT could altematively be a selected constant value, or a predetermined function such as a function of elevation/azimuth, etc., based on information compiled in block 520 over long time periods. Integrity monitoring system 500 further includes ground station integrity information processing means 550 which establishes satellite-specific statistical integrity parameters associated with the resultant satellite-specific differential coπection infoπnation for deriving integrity information associated with each ofthe satellite- specific differential coπection average values C - *** - n ' k determined by the satellite- specific differential coπection processing means 300. As illustrated in Figure 5, discriminator 510 receives as inputs (i) the outputs of the measurement time synchronization processing block 420, and (ii) the output of a YI lc ~ ~fl k differential coπection averaging block 430, namely C ' and C ' , respectively. In turn, the discriminator 510 serves to operate on the aforesaid input values for a determination of a discriminant value Z"' which, in one embodiment ofthe invention, represents the difference between each ofthe receiver-satellite-specific coπection values Λ „ K ~n k C ' and the differential coπection average value C ' . Statistical value processing means 520, in the prefeπed embodiment ofthe invention, is operative for determining receiver-satellite-specific statistical values σ* associated with the coπesponding behavior ofthe receiver-satellite-specific integrity discriminant values Z"' , representative ofthe statistical behavior ofthe integrity discriminant value Z n ' over a selected time interval, for example a 24-hour measurement period, or weeks of monitoring. Statistical value processing means 520 may simply be a mathematical processor for determining the receiver-satellite-specific standard deviation of each ofthe receiver- satellite-specific integrity discriminant values Z n ' over the same time period. In turn, vt lc the statistical integrity discriminant value σ ,' may be employed by both the deselecting integrity system 540 and ground station integrity information processing means 550. In one embodiment ofthe invention, discrete data integrity monitoring means 540 includes detection threshold processing means 544 and comparator 546. Detection threshold processing means 544 serves to calculate a detection threshold value DT based on (i) the output of statistical value processing means 520, and (ii) prescribed probability limits, as will be described later, and mathematically expressed by: DT n ' k =σ n ^Q ' « Comparator 546 is operative for comparing each ofthe integrity discriminant values Z n ' with the detection threshold DT and providing an output Fail [n,k] intended to be an identification ofthe particular satellite vehicle transmitting a satellite information signal received by a particular one ofthe ground station GPS signal receivers which yields a differential coπection value beyond selected limits~i.e., failing the detection threshold test ofthe integrity discriminant value. In turn, the output Fail[n,k] of comparator 546 appropriately provides an input to deselection block 460 illustrated in Figure 4. Refeπing again to Figure 5, ground station integrity information processing means 550 is shown responsive to the statistical integrity discriminant values a n ? and an input K* indicative ofthe number ofthe remaining operative ones ofthe GPS signal receivers 100A, 100B, and lOOC. In turn, ground station system integrity processing means 550 operates on the statistical integrity discriminant values σ^' for determining satellite-specific integrity infoπnation associated with each ofthe satellite-specific differential coπection average values C ' . In one embodiment of the invention, ground station system integrity processor 550 is operative for providing three satellite specific integrity parameters: D fail; σ f a [\_ and σ x wh j ch ^ intended to be utilized by an DGPS airbome flight control system as will be subsequently described. As indicated above, the discrete data integrity monitoring system 540 is intended to provide an indication of which receiver-satellite-specific integrity discriminant values exceed the detection threshold value DT provided by detection threshold processor 544. Comparator 546 is intended to output a receiver-satellite- specific deselection matrix Fail[n,k].which in turn is provided as an input to deselection block 460 as illustrated in Figure 4. Deselection block 460 then provides the receiver- satellite-specific deselection matrix Fail[n,k] as inputs to satellite-specific differential coπection averaging means 430 for eliminating those specific receiver-satellite-specific coπection values 1, which were found to result in failing the detection threshold test from being considered in the determination ofthe differential coπection average value It should be noted that those coπection values C n ' which result in a failure to pass the detection threshold test at one moment~and one satellite constellation configuration, may be acceptable at other moments. However, deselection block 460 may also include software for keeping track of a pattern of behavior of any specific GPS signal receiver which results in differential coπection failures which consistently are identified as failing the detection threshold, i.e., determining if particular GPS signal receivers and/or satellites have continuously produced resultant failures, i.e. Fail [n,k].. In turn, the failing one may be considered for elimination of use of any output therefrom, i.e., "permanently" deselecting such receivers and/or satellite differential coπection infoπnation, indicate a request for service, and/or issue a notice of a DGPS ground station failure. This is so indicated by Fail [n,k]. As illustrated in Figure 5, the deselection matrix may also be provided as inputs to one or more ofthe blocks which comprise the integrity monitoring system 500. Use of deselecting those receiver-satellite measurements and coπespondind data values may be advantageous for minimizing processing time as well as enhance the closed loop processing scheme. That iis, as illustrated in Figure 5, the output of compartor 540, namely the deselection matrix, is dependent upon values which were determined based on presumed valid data. However, after a measurement fails, the average value ofthe differential coπection is affected, and accordingly so is th4 detection threshold as well as the integrity parameters. Thus employment ofthe deselection matrix provides a very robust integrity monitoring system in accordance with the present invention. It should be obvious to those skilled in the art that the three ground station GPS signal receivers provide a fail-safe, fail-operational system for determining the satellite- specific differential coπection information. Whereas, if one ofthe GPS signal receivers should fail, fail-safe operation still would be available pending voter techniques for fail¬ safe, fail-operative operation, the details of which are beyond the scope ofthe present invention. However, a voter system is envisioned as part of a complete fail-safe, fail- operative system in practice ofthe present invention. Illustrated in Figure 6 is an DGPS airbome landing system, similar to that illustrated in Figure 2, but which, in accordance with the present invention, incoφorates differential coπection information and associated integrity information therewith for providing a highly reliable DGPS landing system capable of meeting CAT-I flight control and landing system requirements. As illustrated in Figure 6, the airbome DGPS landing system in accordance with the present invention incoφorates DGPS receiver 252 which serves in substantially the same manner as that described with reference to Figure 2. However, the DGPS airbome landing system in the present invention further includes an integrity monitoring system 600 and a runway lateral-vertical solution matrix transformation processing block 610 for matrix transformation from the usual position solution matrix Sg EE associated with the airbome DGPS signal receiver to a runway lateral/vertical position solution matrix S H y Further, differential coπection information receiver 290' is similar to differential coπection receiver 290 illustrated in Figure 2, but which includes (i) means for receiving the differential coπection values C n derived from the plurality of GPS signal receivers 100A-C, indicated by output data flow line 296 and (ii) means for retrieving integrity information associated with each ofthe satellite-specific coπection values C n indicated by output data flow line 298. Integrity monitoring system 600 includes software which serves as a vertical integrity processor 602 and lateral integrity processor 604. Vertical integrity processor 602 is intended to calculate a vertical integrity limit value, VIL, associated with the vertical deviation between (i) the aircraft position derived from the DGPS solution using the differential coπection information transmitted from the ground station, and (ii) a designated glide path. Lateral integrity block 602 is intended to calculate a lateral integrity limit value, LIL, associated with the lateral deviation between (i) the aircraft position derived from the DGPS solution using the differential coπection infoπnation transmitted from the ground station, and (ii) a designated glide path. The vertical integrity value is intended to indicate the degree of tmst ofthe vertical deviation with a selected probability of being valid, and a lateral integrity value is intended to indicate the degree of t st ofthe lateral deviation with a selected probability of being valid. These probabilities range from [1-10 ] to [1-10 9 ]. Airbome DGPS signal receiver 252, illustrated in Figure 6, like that illustrated in Figure 2, includes pseudo range coπection processing (not shown) responsive to the differential coπection values as outputs on data line 296 from airbome differential coπection infoπnation receiver 290'. In turn, as before, airbome DGPS signal receiver 252 includes a means for determimng the aircraft position A(x,y,z) and position solution matrix SgcEE ■ indicated by numerals 654 and 655, respectively. Runway lateral- vertical transformation processing means 610 is intended to provide the following transformation from S ECEF to S R V having the form: where r, 1, v, and t represent runway, lateral, vertical, and time, respectively. This transformation is commonly known in the art and is the basis for describing the lateral and vertical deviation between the DGPS derived position ofthe aircraft and the designated glide path. Airbome integrity processing means 600 further includes a comparator 606 for comparing the vertical integrity limit value output VIL ofthe vertical integrity processing means 602 with a predetermined maximum vertical integrity limit value identified as VIL [MAX]. Similarly, comparator 608 compares the lateral integrity value output LIL ofthe lateral integrity processing means 604 with a predetermined maximum lateral integrity limit value identified as LIL[MAX]. In turn, these comparators 606 and 608 may provide alarm indicators 625 and 635, respectively. These alarm indicators may be used to audibly or visually alert a pilot, or alert other flight management subsystems (for example, navigation guidance control block 690) for forcibly determining recognition that the differential coπection values and the integrity associated therewith are beyond specified limits for executing a DGPS driven landing. In the embodiement ofthe invention, IOD changes must be handled coπectly to assure that the detection thresholds are not triggered by such events. This can be done by letting GPS signal receiver 100 provides double sets of satellite positions and the difference in the satellite differential coπections using ephemeris data at old and new times. When all receivers (block 100) have received the new ephemeris, a timer is set to initiate the switch to the new ephemeris instantaneously in all K receivers. The timer is set long enough to assure that the airbome user's GPS signal receivers have the new ephemeris available. Utilization of this technique assures that the monitoring scheme described above is realizable. The following mathematical analysis is intended to provide the reader a further understanding ofthe integrity monitoring system 500 in accordance with the present invention. First, it should be noted that each receiver-satellite-specific differential coπection may contain a noise value " w w ' , where k represents the specific receiver. Then: where k=l through K total receivers, except for the differential coπection of a malfunctioning receiver "m" with an offset ε described by: C n>m = C n>m {true} + w n>m + ε In order to detect a malfunctioning receiver result, the average value of all ofthe satellite-specific differential coπection values for all receivers is first established by satellite-specific differential coπection average processing means 430. The satellite- specific differential coπection average may be mathematically expressed by: _ 1 N F N k=l where C n ' is the receiver-satellite-specific differential coπection value at the common synchronization time, namely the output of measurement time synchronization processing block 420. A discriminant may be chosen as a function ofthe satellite-specific differential coπection average value as follows: d[m\ = dp corrm [m\-m p for each ofthe K receivers and N satellite vehicles. In order to meet FAA requirements, it is mandatory or assumed that the discriminant be less than a detection threshold value DT which is determined as a function of a selected false detection rate pf d as follows: Defining 0 ~] as the inverse of Q then: D = σ d Q -^—) , for example if the chosen false detection rate of 10 5 /sample, then p fd = \0 v-5 J => D = 4Λσ d where \d[m]\ > D represents the 1 -sigma standard deviation ofthe discriminant Z n ' , which is determined by statistical integrity discriminant value processing block 520. One simple implementation for statistical integrity discriminant value processor 520 is collecting receiver-satellite-specific discriminant values Z"' , the output of integrity discriminant value processor 540, at a selected sample rate and over a selected time interval, and subsequently determining the 1 -sigma standard deviation ofthe samples. In essence, each of these receiver-satellite-specific 1 -sigma values is thereby related to acquisition ofthe coπesponding satellite information signal transmitted from the same satellite vehicle (n) by the same GPS signal receiver system (k) for a sufficient time to assess the quality ofthe specific receiver with respect to a specific satellite information signal. Detection and identification of a malfunctioning receiver-satellite combination is determined by comparator 546 whenever: ? n,k > DT n,k and comparator provides the indication of those combinations which produce results which fail the comparison, as identified by Fail[n,k]. It should be noted that associated with each ofthe discriminant values Z"' is noise. Noise in the integrity discriminant value Z"' associated with a malfunctioning receiver "m" may be described as a function ofthe noise in each differential coπection value associated with a particular receiver by: From the above expression, it follows that the square ofthe 1 -sigma discriminant standard deviation may be related to the square ofthe 1 -sigma variation in the receiver noise as follows: ' σ < -* = ( ~ X^> σ ) + (— ) Λσ / J w ' ) = —— σ^ ) " A k≠m Λ The standard deviation ofthe noise as mathematically described above is very useful for determining satellite-specific parameters descriptive ofthe accuracy of each ofthe satellite differential coπection values determined by the ground station in accordance with the present invention, more specifically, the satellite-specific differential coπection average values C n . One such parameter is commonly refeπed to as the UDRE (User Differential Rate Eπor @ 99.5% performance ) which describes accuracy under fail free conditions by the following exemplary mathematical relationship: UDRE n ' k = 3.0 - * * τσ%* = 3.0 J— o n J k , and n,k _ \_ K w' ~ ^Yx^ d where σ'l' is the 1 -sigma standard deviation discriminant value evaluated for the satellite specific value of Z n ' for all K receivers and for all N satellites where the observations are desirably taken over at least 24 hours of operation. An enhanced approach in accordance with the present invention is the determination ofthe receiver-satellite-specific 1-sigma values σ'V profiled for satellite elevation and azimuth angles. In other words, a table of "receiver-specific look-up table" may be established for all ofthe satellites based on their elevation and azimuth angles relative to the GPS signal receiver (being substantially the same for those in close proximity to the runway). The look-up table may be determined by a variety of techniques - for example, sampling the discriminant values as before, but also identifying the elevation and azimuth angles at the time ofthe sample. This approach enhances the validity ofthe ground station satellite-specific differential coπection integrity information more fully described below. Therefore, in the exposition which follows, the value for σ n y k may be assumed to be either the usual receiver-satellite-specific standard deviation value, or altematively that which is evaluated particularly for satellite-specific elevation/azimuth position. As indicated earlier, an airbome flight control system requires a DGPS aircraft position to be highly accurate with several integrity requirements. More particularly, any eπors in the lateral and vertical deviation between the aircraft position and a prescribed flight path is of paramount importance for success of DGPS driven precision landings. The following mathematical description sets forth ground station integrity parameters in accordance with the present invention, and is intended for employment in an airbome DGPS integrity monitoring system. The airbome integrity monitoring system is intended to provide a pilot an indication when the overall integrity ofthe DGPS ground station is sufficient to rely upon a DGPS driven landing which meets those requirements for CAT-I landing as required by the FAA. It should be noted that at detection the following holds tme: ±DT n,m = z n,m = KX- (ε + w n,m ) _∑ * w n.k K K k≠m which may be reaπanged as follows: which represents the noise and eπor in a malfunctioning receiver "m" as a function of the detection threshold value DT and the noise in non-malfunctioning receivers. Since the integrity ofthe system is dependent on the differential coπection average value C , it is important to evaluate the worst case eπor in this value as mathematically described by: Kc where K is the number of operative ground station GPS signal receivers, and Wfan is the noise in each averaged differential coπection value under a failure condition mathematically expressed by: σ ∑(σ^) 2 fail (K- ■ l) 2 k≠m Assuming a normal distribution function in non-malfunctioning satellites, the worst case noise in the satellite-specific differential coπection average value C may also be mathematically expressed as a function of the noise in each satellite measurement σ ' w : Finally a weighted value ofthe detection threshold is chosen to be: f a K-\ In one embodiment ofthe present invention, the ground station differential coπection integrity information processing 550 is configured to provide as outputs: D f « ail , a M » , and o» . Now considering an airbome DGPS flight control system, it should be understood that a DGPS driven flight control system makes use of the position solution matrix S E EF output (related to the line of sight measurement matrix, sometimes refeπed to as the H matrix) provided by the DGPS signal receiver 252, but which is transformed into a position solution matrix expressed in terms of a "runway-lateral- vertical coordinate reference frame", identified as S RLV . This latter matrix, is identified by numeral 611, the output of block 610 in Figure 6. The matrix elements have following form: where S j nγ represents the least squares position solution matrix in the runway-lateral- vertical coordinate reference frame by application of a transformation ofthe position solution matrix SgcE/ • ^ snou ld be noted that the position solution matrix is in the form of 4 X N matrix, where one ofthe rows, identified as row "L" represent the lateral coefficients, and another row identified as row "V" represent the vertical coefficients, the other two rows being associated with runway direction, and time. Refeπing again to Figure 6, it is the puφose of vertical integrity processor 602 to evaluate the vertical integrity limit value for the "j th " satellite as mathematically described by: where +(c ab r]+(syrt(° J faι r+(° J ab +(°FTE v r where the value σ p _v γ represents the 1 -sigma standard deviation ofthe "Flight Technical Επor" in the vertical direction. This value is, of course, weather and aiφlane dependent, and typically is an empirically determined value. This assumes the σ"' is the same for all receivers, namely σ" . Altematively, an actual FTΕ_V value may be observed in real time by empirical measurements as opposed to the 1 -sigma value. In this case, the FTE_V value would be eliminated from the just recited expression and added to the preceding expression for vπJ . Similarly, lateral integrity processor 604 is intended to evaluate the lateral integrity limit value for the j th satellite as mathematically described by: LIL J = s S > iL D D "fl J fnaiil + < " y' L Q l (Pmd) where, (σ{) 2 H* n ab ) 2 ]Hs J L ) 2 [(σ' ) Z + (° J ab ) 2 Hσ F TE L T where the value csfpp ι represents the 1 -sigma standard deviation ofthe Flight Technical Eπor in the lateral direction. It too may altematively be eliminated from the just recited expression, and the observed value FTEJL added to the preceding expression for LW . Although not mathematically expressed herein, it should be recognized that the integrity ofthe resultant runway direction or heading (RIL) may be calculated as well in a similar manner and it's impact related to the VIL and the desired glide path angle. In one embodiment ofthe invention, vertical integrity processor 602 is operative for providing as an output the vertical integrity limit value (VIL) maximum of any of the tracked satellite at the cuπent landing system sample time. In turn, this value may be compared with a selected maximum VIL[MAX] by comparator 606. Similarly, lateral integrity processor 604 is operative for providing as an output the lateral integrity limit value (LIL) maximum of any ofthe tracked satellite at the cuπent landing system sample time. In turn, this value may be compared with a selected maximum LIL[MAX] by comparator 608. In turn, whenever either comparators 606 or 608 detects that either the vertical or lateral integrity values exceeded their respective maximum threshold value, comparators 606 and 608 output a selected alarm signal for subsequent system process to notify flight management systems or the pilot directly ofthe alarm situation in order for coπective action to be taken. It should be recognized that the schematic block diagrams particularly illustrated in the accompanying Figures may be configured by employment of many electronic subsystems, computers, and software/firmware implementations, all of which are intended to be within the tme spirit and scope ofthe present invention. In particular the processing means and or designated blocks are intended to perform the indicated functions and/or calculations. With reference to Figure 3, it should be noted that differential coπection processing means 300 is illustrated as receiving as inputs the measurement information from the plurality of GPS receivers lOOA-C. However, it is within the scope ofthe present invention that the differential coπection processing means 210 may be incoφorated as part ofthe GPS receiver, itself. In this situation, the differential coπection processing means 300 would perform the remaining functions as illustrated in Figure 4 except for the raw differential coπection. It should be understood by those skilled in the art the location ofthe processing function is a matter of design choice, all of which are intended to be within the scope of the present invention. Similarly, ground station integrity information processing means 550 may altematively be on the DGPS airbome system illustrated in Figure 6. In this situation, the ground station is operative to transmit the information necessary to obtain the needed satellite-specific integrity information. For example, only the satellite-specific statistical values and the number of operative receivers is what is required to obtain the necessary integrity information as indicated by the above mathematical analysis. Selected discriminant and algorithms have been described herein for obtaining integrity information associated with the satellite-specific differential coπection values. Others, of course, may be applied without departing from the tme spirit and scope ofthe present invention, namely, the method of obtaining integrity information for a landing system where integrity information is associated with a satellite-specific differential coπection value derived by averaging the satellite-specific differential coπection value determined by a plurality of DGPS receivers. Lastly, particular vertical and lateral integrity limit values have been shown with regard to a particular runway later vertical transformation matrix. However, other such transformations and coordinate systems are intended to be within the tme spirit and scope ofthe present invention.

Claims: CLAIMS The embodiments of an invention in which an exclusive property or right is claimed are defined as follows: 1. A differential satellite positioning system ground station for determining GPS differential coπection infoπnation with integrity monitoring, said differential satellite positioning system ground station comprising: a plurality of GPS signal receivers, each at a known geocentric position, each operative for receiving a satellite-specific information signal transmitted from each of selected ones of a plurality of satellite vehicles, and each operative for deriving, from each of said received satellite-specific information signals from said satellite vehicles, receiver-satellite-specific GPS signal information; first differential coπection processing means for deriving receiver-satellite- specific GPS differential coπection information associated with each of said satellite-specific information signals, where said receiver-satellite specific differential coπection information is a function of said receiver- satellite-specific GPS signal information determined by a specific one of said GPS signal receivers and said known position of said GPS signal receivers for the same one of said GPS signal receivers; second differential coπection processing means operative for determining satellite-specific GPS differential coπection information, associated with each of said satellite-specific information signals, as a function of said receiver-satellite-specific GPS differential coπection infoπnation determined by more than one of said plurality of GPS signal receivers and associated with the same one of said satellite-specific information signals; ground station integrity monitoring means for determining satellite-specific integrity information associated with each of said satellite-specific information signals, as a function of said receiver-satellite-specific GPS differential coπection information and said satellite-specific GPS differential coπection infoπnation, where said satellite-specific integrity information is descriptive ofthe integrity of said satellite-specific GPS differential coπection information. 2. The differential satellite positioning system ground station of claim 1 wherein: said first differential coπection processing means is operative for determining a receiver-satellite-specific GPS differential correction value, associated with each of said satellite-specific information signals and each of said plurality GPS signal receivers, as a function of said receiver-satellite- specific GPS signal information, and said known position ofthe coπesponding one of said plurality of GPS signal receivers; said second differential coπection processing means includes means for providing a satellite-specific statistical GPS differential coπection value, associated with each of said received satellite-specific information signals, as an average of said receiver-satellite-specific differential coπection values associated with the same satellite information signal and derived from said receiver-satellite-specific GPS signal information from more than one of said plurality of GPS signal receivers; and said ground station integrity monitoring means includes, discriminator means operative for determining, for each of said receiver- satellite-specific GPS differential coπection values, a receiver- satellite-specific integrity discriminant value as a selected function of coπesponding ones of said satellite-specific statistical GPS differential coπection values and said receiver-satellite- specific differential coπection values associated with the same satellite-specific information signal and coπesponding one of said plurality of GPS signal receivers, a statistical value processing means operative for determimng a receiver- satellite-specific statistical value associated with each of said receiver-satellite-specific integrity discriminant values, as a function of coπesponding ones of said receiver-satellite-specific integrity discriminant values, and which is descriptive ofthe behavior of said receiver-satellite-specific integrity discriminant values sampled over a selected time period. 3. The differential satellite positioning system ground station of claim 2 wherein said ground station integrity monitoring means determines said satellite- specific integrity information as a function of said receiver-satellite-specific statistical integrity discriminant values associated with more than one of said plurality of GPS signal receivers and associated with the same one of said satellite-specific information signals. 4. The differential satellite positioning system ground station of claim 2 wherein said integrity monitoring means further includes deselecting means for identifying those of said receiver-satellite-specific integrity discriminant values which exceed a selected detection threshold. 5. The differential satellite positioning system ground station of claim 4 wherein said deselecting monitoring means includes: means for determining a detection threshold value as a function of said receiver- satellite-specific statistical integrity discriminate values; and comparator means for comparing each of said integrity discriminate values with said detection threshold value, and providing an identification of which receive-satellite-specific ones of said receiver-satellite-specific discriminant values exceeds said detection threshold value. 6. The differential satellite positioning system ground station of claim 1 further comprising means for communicating said satellite-specific differential ' coπection information and coπesponding satellite-specific integrity information to a remote differential satellite positioning system user. 7. The differential satellite positioning system ground station of claim 1 wherein said differential coπection processing means includes means for determining said receiver-satellite-specific differential coπection values synchronized to a common measurement time and with receiver-specific clock bias removed. 8. The differential satellite positioning system ground station of claim 5 wherein each of said satellite-specific differential coπection values is a function of said receiver-satellite-specific differential coπection values excluding those receiver- satellite-specific ones having coπesponding receiver-satellite-specific discriminant values which were identified by said comparator means. 9. A differential satellite positioning system comprising: ground station for determining GPS differential coπection information with integrity monitoring including, a plurality of GPS signal receivers, each at a known geocentric position, each operative for receiving a satellite-specific information signal transmitted from each of selected ones of a plurality of satellite vehicles, and each operative for deriving, from each of said received satellite-specific information signals from said satellite vehicles, receiver-satellite-specific GPS signal information; first differential coπection processing means for deriving receiver- satellite-specific GPS differential coπection information associated with each of said satellite-specific information signals, where said receiver-satellite specific differential coπection infoπnation is a function of said receiver-satellite-specific GPS signal information determined by a specific one of said GPS signal receivers and said known position of said GPS signal receivers for the same one of said GPS signal receivers; second differential coπection processing means operative for determining satellite-specific GPS differential coπection infoπnation, associated with each of said satellite-specific information signals, as a function of said receiver-satellite- specific GPS differential coπection information determined by more than one of said plurality of GPS signal receivers and associated with the same one of said satellite-specific information signals; ground station integrity monitoring means for determining satellite- specific integrity information associated with each of said satellite-specific information signals, as a function of said receiver-satellite-specific GPS differential coπection infoπnation and said satellite-specific GPS differential coπection information, where said satellite-specific integrity information is descriptive of the integrity of said satellite-specific GPS differential coπection infoπnation means for transmitting said satellite-specific differential coπection information and said satellite-specific integrity information to a remote differential satellite positioning system; and a remote differential satellite positioning system including, means for receiving said satellite-specific differential coπection infoπnation and said satellite-specific integrity information, GPS signal receiver operative for (i) receiving a satellite-specific information signal transmitted from each of selected ones of a plurality of satellite vehicles, and operative for deriving from each of said satellite-specific information signals from said satellite vehicles user-satellite-specific GPS signal information,(ii) deriving the user-position of said remote GPS signal receiver as a function of said user-satellite-specific GPS signal information coπected by said satellite-specific GPS differential coπection information, and (iii) determining a position solution matrix representative ofthe line of sight vectors between the derived user-position and each of said satellite vehicles, and integrity monitoring means for determining the integrity of said user position as a function of said position solution matrix and said satellite-specific integrity infoπnation. 10. A differential satellite positioning system ground station for determining satellite-specific GPS differential coπection information with coπesponding satellite- specific integrity information, said differential satellite positioning system ground station comprising: a plurality of GPS signal receivers, each at a known geocentric position, each operative for receiving a satellite-specific information signal transmitted from each of selected ones of a plurality of satellite vehicles, and each operative for deriving, from each of said received satellite-specific information signals from said satellite vehicles, receiver-satellite-specific GPS signal information; differential coπection processing means for separately deriving a receiver- satellite-specific GPS differential coπection value for each GPS signal receiver, and associated with each of said satellite-specific information signals, as a function of said receiver-satellite-specific GPS signal information determined by each of said plurality of GPS signal receivers; means for determining a satellite-specific GPS differential coπection value, associated with each of said satellite-specific information signals, as an averaging function of said receiver-satellite-specific GPS differential coπection values determined by more than one of said plurality of GPS signal receivers and associated with the same one of said satellite- specific information signals; discriminator means for determining a receiver-satellite-specific discriminant value associated with each satellite-specific information signal and specific one of said GPS signal receivers , where each receiver-satellite- specific discriminant value is a function ofthe difference between coπesponding ones of said receiver-satellite-specific GPS differential coπection value and said satellite-specific GPS differential coπection value associated with the same one of said satellite-specific information signals and specific one of said GPS signal receivers statistical value processing means for determining a receiver-satellite-specific statistical value descriptive ofthe coπesponding behavior of each of said receiver-satellite-specific discriminant values detection threshold processing means operative for calculating at least a satellite- specific detection threshold value as a function of a selected false detection probability and coπesponding receiver-satellite-specific statistical value; and ground station integrity monitoring means for determining satellite-specific integrity information, associated with each of said satellite-specific information signals and coπesponding satellite-specific GPS differential coπection values, as a function of said receiver-satellite-specific statistical value, said satellite-specific detection threshold value, and the number of operating GPS signal receivers. 11. The differential satellite positioning system ground station of claim 10 where said receiver-satellite-specific statistical value is descriptive of the behavior of said receiver-satellite-specific integrity discriminant values sampled over a selected time period. 12. A differential satellite positioning system ground station for determining GPS differential coπection infoπnation with integrity monitoring, said differential satellite positioning system ground station comprising: a plurality of GPS signal receivers, each at a known geocentric position, each operative for receiving a satellite-specific information signal transmitted from each of selected ones of a plurality of satellite vehicles, and each operative for deriving from each of said received satellite-specific information signals from said satellite vehicles receiver-satellite-specific GPS signal information; differential coπection processing means for determining a receiver-satellite- specific differential coπection value associated with each of said selected satellite information signals; averaging means for determining a satellite-specific average value of said receiver-satellite-specific differential coπection values associated with a mutually exclusive one of said satellite information signals; discriminator means responsive to receiver-satellite-specific differential coπection values and said satellite-specific differential coπection average values for determining a receiver-satellite-specific discriminant value; means for determining a statistical value representative ofthe behavior of said discriminant values; and integrity monitoring means responsive to said discriminant values and said statistical values for deriving selected integrity information as one or more selected functions thereof.

Inventors: BRENNER MATS A, BRENNER, MATS, A.
Applicants: HONEYWELL INC, HONEYWELL INC.
Application year: 1996
Main classification: G01S5
IPC codes: G01C21, G01S1, G01S5