Classifications

• • G—PHYSICS
  • G01—MEASURING; TESTING
  • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
  • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
  • G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
  • G01S19/13—Receivers
  • G01S19/20—Integrity monitoring, fault detection or fault isolation of space segment
• • G—PHYSICS
  • G01—MEASURING; TESTING
  • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
  • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
  • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
  • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
  • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
  • G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments

Definitions

• the present invention relates generally to aircraft navigation and more particularly to a system employing a Global Positioning System(GPS) and an Inertial
• GPS systems have been used to determine aircraft position by receiving signals from a plurality of satellites.
• the signals each have information as to the position of the satellites and the time of transmission so that the GPS receiver, on the aircraft, can calculate its own position. Since there are four variables (position in 3 axes and time), signals from at least 4 satellites are necessary for a determination of receiver position. If there are at least five satellites having good geometry each subset of four signals can be used for positioning and they can be compared with each other to determine if one of the signals is in error (Fail safe).
• the present invention addresses these and other needs by:
• the IRU acceleration signal is compared to the GPS acceleration signal along the predetermined axis. When there is an error greater than a predetermined amount, the faulty satellite signal is identified and can be eliminated from the calculations. Thus only four satellites are necessary for continued operation.
• a signal drift which may be considered as a ramp output
• the initial change at the beginning of the ramp causes an acceleration transient which is detected by the acceleration monitor so that early identification of this condition is made.
• FIG. 1 shows a schematic block diagram of the system utilizing the present invention
• FIG. 2 is a block diagram showing the operation of the acceleration monitor of the present invention.
• FIGS. 3 A and 3B are graphs showing how the new satellite drift monitor of the above mentioned copending application operates.
• GPS receiver 12 is of a type well known in the art which operates to receive signals from a plurality of satellites, e.g. SI - S6 in Figure 1, indicative of their positions and of the time of transmission. While 6 satellites have been shown, it will be understood, that more or less than 6 satellites may be in view of the GPS receiver 12 and that the number may change over a period of time. GPS receiver 12 operates on the signals from satellites SI - S6 to determine the position of the receiver and to produce the GPS code based pseudo range measurements , pr, to each satellite. Also, by tracking and counting the cycles (including fractions of cycles) in the carrier from each satellite, the GPS receiver 12 provides the accumulated carrier count, pc, which can be used to obtain the change of pseudo range over a time ⁇ t according to the equation:
• the change of the pseudo range obtained by differencing the carrier count is accurate to within centimeters ( 0.01 meter ) while the pseudo range change formed by differencing the pseudo range measurements is accurate to within 2 to 5 meters.
• the code based change in pseudo range ⁇ pr , the carrier based change in pseudo range, ⁇ pc, or a combination thereof are used in the present invention as will be described below.
• the pr and pc signals are provided as outputs represented by a path 16 from GPS receiver 12 and, via paths 17 and 18 to a standard RAIM shown in box 20 which also receives a pressure altitude signal from an altitude sensing means 22 via a path 24.
• a deselect signal is presented from RAIM 20 as an output represented by a path 26. This occurs if there are at least 6 satellites in view with good geometry, in which case, groups of five satellite signals can be compared so as to determine which one, if any, of them is faulty. If one is so identified, then this information is passed (via path 26 and a path 28) to a satellite selection function represented by a box 30 which also receives the GPS pr and pc from path 16. The satellite selection function 30 eliminates the faulty satellite signal and passes the remaining valid signals as an output represented by a path 32 to a Position function represented by box 34.
• Position function 34 also receives a pressure altitude signal from an Altitude Sensing means 22 via path 24 and produces an output indicative of the aircraft position (as represented by a path 36) to downstream equipment such as indicators or flight management systems (not shown). If there are only 5 satellites in view with good geometry, then groups of four satellite signals can still be compared against at least one other satellite signal so as to determine if any of them are faulty but since the faulty satellite cannot be identified no de-selection can be performed. In such a case the downstream equipment and the pilot will be notified that the GPS signals should not be used (this signal is not shown in figure 1).
• the signal output represented by path 16 is provided via paths 17 and 18 to an acceleration monitor as represented by a box 44.
• the acceleration monitor 44 also receives an input from the inertial reference unit 14 (which is also well known in the art and comprises a plurality of gyros and accelerometers) that produces outputs as represented by a line 46 indicative of h (altitude), C, (attitude matrix), D (latitude, longitude, wander angle matrix), v (velocity) and ⁇ v (change of velocity).
• the inertial reference unit 14 which is also well known in the art and comprises a plurality of gyros and accelerometers
• a line 46 indicative of h (altitude), C, (attitude matrix), D (latitude, longitude, wander angle matrix), v (velocity) and ⁇ v (change of velocity).
• a minimum of 3 gyros and 3 accelerometers are employed, but to ensure fail safe or fail operational operation and high reliability, two or three redundant systems are preferably employed.
• the ⁇ v signal from path 46 and the pr, pc signals from paths 16, 17 and 18 are used by the Acceleration Monitor, 44, of the present invention by an operation which will be better understood with reference to Figure 2.
• the Acceleration Monitor 44 is shown receiving the inertial signals h, C , D v, and ⁇ v over a path shown by arrow 46 which corresponds to figure 1 and receiving the GPS input, pr and pc, over a path shown by arrow 18 which corresponds to Figure 1.
• the inertial signal ⁇ v consists of filtered velocity increments at a 10 - 60 Hz rate which are provided to a function box 54 labeled "Transform to L frame".
• Function box 54 also receives the attitude input C, shown by arrow 46 and transforms the filtered velocity increments to the local vertical frame, L, and outputs the transformed increments to a function represented by a box 58 labeled "Form 2nd 1 Hz Time Difference".
• box 58 "double position difference signals” (i.e. signals representing the difference in the change of position in the current ⁇ t interval and the change of position in the previous ⁇ t interval) along each axis are formed by integrating the inertial acceleration (i.e. the transformed high rate velocity increments).
• the output of the function 58 reflects an inertially measured acceleration which contains earth rotation induced acceleration components that will not be present in the acceleration derived from the GPS signal (to be explained below).
• the output of function 58 is provided as input to a function represented by a box 60 labeled "Remove Earth Rotation Induced Acceleration” where the undesired earth rotation components are removed.
• the inertially based reference signal from function 60 is presented to a function represented by a box 62 labeled "Project Along Line Of Sight” which also receives a signal, LOS ( unit vector along the line of sight to the satellite ) shown by arrow 64 and the output of function 62 is an inertial double position difference projected along the line of sight to the satellite.
• This signal is presented to a function represented by box 66 labeled "Discriminator & time removal transformation" where the acceleration monitor discriminator is formed.
• the pseudo range measurements pr and accumulated carrier cycle counts pc for all tracked satellites are presented from path 18 to a function represented by a box 70 labeled "Form Range Difference" which also receives an initial GPS position, r init , shown on an input represented by arrow 72 which is the GPS position at the time of initialization and received from the position function 34 in figure 1.
• the signal, pc consists of accumulated carrier cycles (each cycle corresponds to about 0.19 m position change in the usual satellite signal) and the signal pr are code based pseudo range measurements.
• the signals pr , pc or any combination thereof include the motion of the satellite and function 70 operates to remove the satellite motion component and provide the result to a function represented by a box 74 labeled "Form 2nd 1 Hz Time Difference".
• the double difference signal i.e. the difference in the change of cycle count ( or smoothed pseudo range ) in the current ⁇ t interval and the change of cycle count ( or smoothed pseudo range ) in the previous ⁇ t interval
• function 76 an output which represents the acceleration of the GPS receiver (but which also contains components that relate to the change in the line of sight vector at the current position and the line of sight vector at the initial reference position) is provided to a function represented by a box 76 labeled "Line of Sight Compensation".
• function 76 is one where the components related to the line of sight are removed.
• the output from function 76 is presented to a function represented by a box 78 labeled
• the cycle count includes the motion of the satellite and function 78 operates to remove the satellite motion acceleration component.
• the final result is a GPS signal based double position difference along the line of sight and this is provided to the function represented by box 66.
• the "Discriminator & Time Removal Transformation” function 66 operates on the GPS and Inertial Reference acceleration signals from functions 78 and 62 respectively to form a discriminator by differencing the two signals. This function also subtracts an average (over all of the satellites) of all discriminators from each satellite specific discriminator thereby eliminating the receiver clock offset.
• This signal is presented to a function represented by a box 80 where the discriminator output is averaged over time and compared to a fixed threshold value to produce a deselected signal on a line 82 which is used to deselect any satellite whose acceleration exceeds the threshold. It should be noted that all of the functions performed by function boxes described above can be performed by a computer program with each function being readily programmable by one having ordinary skill in the programming art.
• the deselect signal on line 82 of figure 2 is shown being provided via lines 26 and 28 to the Select Satellite function 30 and the deselected satellite signal will not proceed to the Positioning function 34.
• An Acceptance Monitor 86 is shown in figure 1 receiving the GPS signal over paths 16, 17 and 18 and further receiving a pressure altitude signal from the Altitude box 22 over path 24.
• the Acceptance Monitor 86 is used in the art to detect GPS signals which are obviously incorrect because, for example the satellite pseudo range is far out of reasonable bounds. When such an erroneous satellite is received, the
• Acceptance Monitor 86 produces a deselect signal to path 26 and via path 28 to the Select Satellite function 30 which then operates to prevent the erroneous signals from being used by the Positioning function 34.
• a new satellite drift monitor or "Z RAIM” is shown as a box 92 which receives the GPS signals over paths 16, 17 and 18 and the pressure altitude signal over path 34.
• Z RAIM 92 provides an output indicative of the Horizontal Protection Limit, HPL, to the pilot or to downstream aircraft equipment such as the Flight Management System (not shown) over a path 95
• the new satellite drift monitor or Z RAIM 92 of the present invention (which is claimed in the above referred to copending application) operates to determine if a drift has already begun when a signal from a new satellite is first acquired. This is explained with reference to Figures 3 A and 3B.
• the RAIM discriminator Assuming that there are N satellites with good geometry, the RAIM discriminator, as known in the art for the nth satellite, is formed by the equation:
• b is a well known satellite geometry dependent coefficient as seen in the above k referred to publication
• N is the number of satellites
• k indicates the kth satellite
• ⁇ pr k is the difference between the measured pseudo range (or smoothed pseudo range) and the predicted pseudo range for the kth satellite. Due to selective availability (SA, a deliberate noise signal superimposed on the output of the GPS by the DOD), the discriminator will vary as seen in Figure 3 A
• the new satellite error bound, ⁇ , defined to have a predetermined confidence level of l-p m(j which is typically 99.9%, , is defined by the equation:
• the discriminator will cross zero or reach a minimum value from time to time i.e. at least every 5 minutes. As the discriminator crosses zero or reaches the minimum absolute value
• t is the time since the satellite was first received.
• the satellite error limit, SEL is the minimum of the current satellite error bound ⁇ and the previous satellite error limit with the estimated l-p md drift error bound added.
• SEL(t) min ( ⁇ , SEL (t - ⁇ t) + r ⁇ t) (7)
• ⁇ t is the time step.
• HPL a required output in avionics equipment
• K md is the statistical sigma number corresponding to the missed detection
• the Acceleration Monitor 44 of the present invention should detect an acceleration. To achieve fail operation capability, a de-selection of a new satellite via paths 22, 23 and function 24 in figure 1 should be performed if the rate, r, exceeds a predetermined time dependent threshold. Only newly acquired satellites need to be monitored this way since satellites that are already used for positioning, are being monitored by the Acceleration Monitor 44.

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• Engineering & Computer Science (AREA)
• Radar, Positioning & Navigation (AREA)
• Remote Sensing (AREA)
• Physics & Mathematics (AREA)
• General Physics & Mathematics (AREA)
• Automation & Control Theory (AREA)
• Computer Security & Cryptography (AREA)
• Computer Networks & Wireless Communication (AREA)
• Position Fixing By Use Of Radio Waves (AREA)
• Navigation (AREA)
• Gyroscopes (AREA)

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• 1998
  • 1998-07-17 US US09/118,046 patent/US5969672A/en not_active Expired - Lifetime
• 1999
  • 1999-06-25 CA CA002337876A patent/CA2337876C/en not_active Expired - Fee Related
  • 1999-06-25 WO PCT/US1999/014282 patent/WO2000004402A1/en not_active Ceased
  • 1999-06-25 JP JP2000560468A patent/JP4446604B2/ja not_active Expired - Fee Related
  • 1999-06-25 DE DE69904187T patent/DE69904187T2/de not_active Expired - Lifetime
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