US 20020120400 A1 Abstract An improved fully-coupled vehicle positioning method and system with differential GPS can substantially solve the problems encountered in either the global positioning system-only or the inertial navigation system-only, such as loss of global positioning satellite signal, sensitivity to jamming and spoofing, and an inertial solution's drift over time. In the present invention, the velocity and acceleration from an inertial navigation processor of the integrated GPS/INS system are used to aid the code and carrier phase tracking of the global positioning system satellite signals, so as to enhance the performance of the global positioning and inertial integration system, even in heavy jamming and high dynamic environments. To improve the accuracy of the integrated GPS/INS navigation system, phase measurements are used and the idea of the differential GPS is employed. A master-slave relative positioning scheme is invented and is effective for high accuracy formation driving and flight.
Claims(3) 1. An improved fully-coupled vehicle positioning system, comprising:
a global positioning system (GPS) processor for providing GPS measurements including pseudorange, carrier phase, and Doppler shift, for a slave system; a data link for receiving GPS-like signal from a master system, where said GPS-like signal is a frequency shift GPS signal and generated by said master system; an inertial measurement unit (IMU) for providing inertial measurements including body angular rates and specific forces; a central navigation processor, which are connected with said GPS processor, said IMU and said data link, comprising an inertial navigation system (INS) processor, a Kalman filter, a new satellites/cycle slips detection module, and an on-the-fly ambiguity resolution module; and an input/output (I/O) interface connected to said central navigation processor; a GPS antenna of said master system for receiving GPS signal; a frequency mixer of said master system for shifting carrier frequency of said GPS signal received from said GPS antenna to generate said GPS-like signal; a data link of said master system for transmitting said GPS-like signal; wherein said GPS measurements from said GPS processor and GPS-like signal from said data link are passed to said central navigation processor and said inertial measurements are injected into said inertial navigation system (INS) processor; wherein said GPS-like signal is processed by said central navigation processor to derive GPS measurements; wherein an output of said INS processor and said GPS measurements are blended in said Kalman filter; an output of said Kalman filter is fed back to said INS processor to correct an INS navigation solution, which is then output from said central navigation processor to said I/O interface; wherein said INS processor provides velocity and acceleration data injected into said GPS processor to aid code and carrier phase tracking of GPS satellite signals; wherein an output of said GPS processor, an output of said data link, and an output of said INS processor are injected into a new satellites/cycle slips detection module to test the occurrence of new satellites and cycle slips, wherein as said new satellites/cycle slips detection module is on, said on-the-fly ambiguity resolution module is activated to resolve global positioning system satellite signal carrier phase integer ambiguities; wherein said on-the-fly ambiguity resolution module outputs the integer ambiguities into said Kalman filter to further improve positioning accuracy, and said INS processor outputs navigation data to said I/O interface. 2. An improved fully-coupled vehicle positioning system, as recited in a RF (radio frequency) unit for converting the RF GPS signals from a plurality of GPS antennas to a plurality of digital base band GPS signals; a correlation and tracking loops for receiving said digital base band GPS signals of said GPS antenna to perform the tracking processing of the GPS signals and to derive GPS carrier phase, psuedorange, and range rate measurements; a satellite and antenna selection module for choosing a GPS antenna with a maximum number of GPS satellites in view and performing the carrier phase difference processing among said GPS antenna to derive the carrier phase difference measurements of said GPS antenna, wherein said GPS carrier phase, psuedorange, and range rate measurements of the chosen GPS antenna with maximum number of GPS satellites in view are sent to said central navigation processor to perform the fully-coupled GPS/IMU processing; an attitude determination processor for receiving the carrier phase difference measurements of said GPS antenna to derive the GPS attitude measurement. 3. An improved fully-coupled vehicle positioning method comprises steps of:
(a) receiving global positioning system raw measurements, including pseudorange, carrier phase, and Doppler shift; (b) receiving a GPS-like signal from a data link; (c) sending said GPS raw measurements to a central navigation processor from a GPS processor and said data link; (d) receiving a plurality of inertial measurements including body angular rates and specific forces from an inertial measurement unit (IMU); (e) computing an inertial navigation solution which are position, velocity, acceleration, and attitude of a vehicle by sending said inertial measurements from said IMU to an inertial navigation system (INS) processor of said central navigation processor for (f) fixing integer ambiguities based on testing the occurrence of new satellites or cycle slips using said GPS rover raw measurements from said GPS processor, GPS reference raw measurements, position, and velocity from said data link, and said inertial navigation solution from said INS processor and send the integer ambiguities to a Kalman filter; (g) blending an inertial navigation solution derived from said INS processor and said GPS raw measurements from said GPS processor and said data link in said Kalman filter to derive INS corrections and GPS corrections; (h) feeding back said INS corrections from said Kalman filter to said INS processor to correct said inertial navigation solution; and (i) sending said inertial navigation solution from said INS processor to an I/O interface, so as to provide navigation data for an on-board avionics system. Description [0001] This is an improved non-provisional application of a previous allowed non-provisional application, application number 09/246883, filed on Feb. 8, 1999, and another non-provisional application, application number 09/661,587, filed on Sep. 14, 2000. [0002] The present invention relates generally to a global positioning system and inertial measurement unit (GPS/IMUT) integrated positioning and navigation method and system, and more particularly to an improved fully-coupled integration method and system of the global positioning system (GPS) receiver and the inertial measurement unit (IMU), which allows the mutual aiding operation of the GPS receiver and the inertial navigation system (INS) at an advanced level with features of inertial aiding global positioning system satellite signal tracking, fuzzy logic for attitude determination, master-slave relative positioning, robust attitude determination, and on-the-fly resolution of GPS carrier phase integer ambiguities and real-time positioning in the differential GPS mode. [0003] The GPS user equipment, which comprises an antenna, a signal processing unit, and associated electronics and displays, receives the signals from the GPS satellites to obtain position, velocity, and time solutions. There are two types of GPS observables: code pseudoranges and carrier phases. Phase measurements are based on two L-band carrier frequencies. One is the L [0004] Because of the navigation message transmitted by the GPS satellites, the positions and velocities of the GPS satellites can be computed. Therefore, the propagating time of a GPS signal can be determined. Since the signal travels at the speed of light, the user can calculate the geometrical range to the satellite. In this way, the code pseudorange measurements can be determined and is degraded by errors, such as ephemeris errors, user and satellite clock biases (including selective availability (SA)), atmospheric effects (ionosphere and troposphere), and measurement noise (receiver error and random noise). These errors not only affect pseudorange measurements but phase measurements. The most obvious difference between both measurements is the measurement error. For phase measurements, the measurement noise is of the order of a few millimeters and for pseudorange measurements, the measurement noise is accurate to about 30 centimeters (for the P code) or 3 meters (for the C/A code). [0005] The Global Positioning System, GPS, contains a number of error sources: the signal propagation errors, satellites errors, and the selective availability. The user range error (URE) is the resultant ranging error along the line-of-sight between the user and the global positioning system satellite. Global positioning system errors tend to be relatively constant (on average) over time, thus giving the global positioning system long-term error stability. However, the signals of the global positioning system may be intentionally or unintentionally jammed or spoofed, or the global positioning system (GPS) receiver antenna may be obscured during vehicle attitude maneuvering, and the global positioning system signals may be lost when the signal-to-noise ratio is low and the vehicle is undergoing highly dynamic maneuvers. [0006] In addition to the unavoidable errors (such as ionospheric delay, tropospheric delay, clock biases, and measurement errors) and the intentional error (such as SA), the GPS measurements (pseudorange and phase) may also be affected by the environment surrounding a GPS user antenna. Like the multipath effect, because of an object nearby the user antenna, the antenna receives not only a direct signal from a GPS satellite but also a second or more reflected or diffracted signals from the object. For a highly dynamic vehicle, the onboard GPS receiver may lose the lock of a GPS signal because the signal-to-noise ratio (SNR) is low or the GPS signal is blocked by the body of its own vehicle. [0007] Typically, the navigation solution is estimated by using the pseudorange measurements. Since the satellite clock biases are provided by the navigation message, for three-dimensional position determination, in addition to the three unknowns in position, the receiver (user) clock bias also needs to be estimated, i.e., there are four unknowns for the navigation solution. As a result, for a stand-alone receiver, the position determination usually needs a minimum of four visible GPS satellites, and the estimated position is accurate to about 100 meters with SA on. In order to improve the accuracy of the estimated position, the phase measurements will be used. Also, to eliminate the most of SA and other common errors (for example, receiver and satellite clock biases), the differential GPS will be employed. As a result, the accuracy of the estimated position is of the order of a few centimeters. However, to achieve the centimeter accuracy, one of the key steps is to resolve carrier phase integer ambiguities. [0008] An inertial navigation system (INS) comprises an onboard inertial measurement unit (IMU), a processor, and embedded navigation software(s), where the components of the IMU include the inertial sensors (accelerometers and gyros) and the associated hardware and electronics. Based on measurements of vehicle specific forces and rotation rates obtained from onboard inertial sensors, the positioning solution is obtained by numerically solving Newton's equations of motion. [0009] The inertial navigation system is, in general, classified as a gimbaled configuration and a strapdown configuration. For a gimbaled inertial navigation system, the accelerometers and gyros are mounted on a gimbaled platform to isolate the sensors from the rotations of the vehicle and then to keep the measurements and navigation calculations in a stabilized navigation coordinate frame. Generally, the motion of the vehicle can be expressed in several navigation frames of reference, such as earth centered inertial (ECI), earth-centered earth-fixed (ECEF), locally level with axes in the directions of north-east-down (NED), and locally level with a wander azimuth. For a strapdown inertial navigation system, the inertial sensors are rigidly mounted to the vehicle body frame. In order to perform the navigation computation in the stabilized navigation frame, a coordinate frame transformation matrix is used to transform the acceleration and rotation measurements from the body frame to one of the navigation frames. [0010] In general, the measurements from the gimbaled inertial navigation system are more accurate than the ones from the strapdown inertial navigation system. And, the gimbaled inertial navigation system is easier in calibration than the strapdown inertial navigation system. However, the strapdown inertial navigation systems are more suitable for higher dynamic conditions (such as high turn rate maneuvers) which can stress inertial sensor performance. Also, with the availability of modem gyros and accelerometers, the strapdown inertial navigation systems become the predominant mechanization due to their low cost and reliability. [0011] Inertial navigation systems, in principle, permit pure autonomous operation and output continuous position, velocity, and attitude data of the vehicle after initializing the starting position and initiating an alignment procedure. In addition to autonomous operation, other advantages of an inertial navigation system include the full navigation solution and wide bandwidth. However, an inertial navigation system is expensive and is degraded with drift in output (position and velocity) over an extended period of time. It means that the position and velocity errors increase with time. This error propagation characteristic is primarily caused by, such as, gyro drift, accelerometer bias, misalignment, gravity disturbance, initial position and velocity errors, and scale factor errors. [0012] Under the requirements, such as low cost, high accuracy, continuous output, high degree of resistance to jamming, and high dynamics, the stand-alone INS and stand-alone GPS have difficulties to perform properly. Therefore, to decrease or diminish the drawbacks for each system (INS and GPS), the integration of both systems is one of the ways to achieve the above requirements. In general, there are three conventional approaches for integrating the GPS and INS. The first approach is to reset directly the INS with the GPS-derived position and velocity. The second approach is the cascaded integration where the GPS-derived position and velocity are used as the measurements in an integration Kalman filter. The third approach is to use an extended Kalman filter which processes the GPS raw pseudorange and delta range measurements to provide optimal error estimates of navigation parameters, such as the inertial navigation system, inertial sensor errors, and the global positioning system receiver clock offset. [0013] However, there are some shortcomings of the above existing integration approaches and they are summarized as follows: [0014] 1. In the conventional global positioning system and inertial navigation system integration approaches, only position and velocity from the output of the GPS receiver or the GPS raw pseudorange and delta range measurements are used. However, the GPS raw phase measurements haven't been used for an integration solution, although the phase measurements are accurate to a few millimeters in contrast to 30 centimeters for the P code pseudorange or 3 meters for the C/A code pseudorange in the presence of measurement noise. [0015] 2. There is a significant impediment to the aiding of the global positioning system signal tracking loops with an inertial navigation system. It is that the aiding causes the potential instability of the conventional global positioning system and inertial navigation integration system because of a positive feedback signal loop in the integrated global positioning and inertial system. As a result, the degradation in accuracy of the inertial aiding data increases the signal tracking errors. And, the increased tracking errors are fed back into the inertial system. This may cause further degradation of the inertial system because the measurements may severely affect the Kalman filter, which is well tuned for a low accuracy inertial navigation system. [0016] 3. The inertial sensors in the conventional tightly-coupled GPS and inertial integration system can not provide the high accuracy in velocity. Therefore, the aiding of a carrier phase tracking loop can not execute properly due to the need for high accuracy of the external input velocity. [0017] An objective of the present invention is to use the velocity and acceleration from an inertial navigation processor, which are corrected by a Kalman filter, as the aiding of the code and carrier phase tracking of the GPS satellite signals so as to enhance the performance of the GPS/INS, even in heavy jamming and high dynamic environments, and to improve the accuracy of the receiver position and velocity by using differential GPS. To accurately determine the receiver position and velocity at the centimeter level, the GPS phase measurements will be used and the differential GPS will be employed. In this invention, a new process (OTF (on-the-fly) technique) is disclosed to resolve the integer ambiguities on the fly and estimate the receiver position in real time. The results of GPS estimates will increase the accuracy of the inertial navigation system and therefore enhance the capability of the GPS tracking loop. [0018] Another objective of the present invention is that the self-contained INS complements the GPS as the GPS receiver loses lock of the GPS signals. Once the GPS receiver regains the signals and then estimates the receiver position and velocity, the output (position and velocity) of the GPS receiver is used to correct the position and velocity of the INS that have drifted. [0019] Another objective of the present invention is that a data link is used to receive the data, such as position, velocity, and raw measurements, from a reference site in addition to a GPS receiver to collect the raw measurements for a rover site. Using the differential GPS and phase measurements, the accuracy of the GPS positioning is of the order of centimeter level after fixing the integer ambiguities, and, as a result, the integrated GPS/INS is applicable in high accuracy positioning. [0020] Another objective is to use fuzzy logic for multi-antenna GPS attitude determination, where false GPS measurements are isolated to enhance the robustness of the attitude determination system. [0021] Another objective is to use a master-slave scheme for relative positioning using GPS which is difficult to intercept indirectly. [0022] A further objective of the present invention is that the inertial navigation system can aid the resolution of the GPS carrier phase integer ambiguities by providing more accurate position information. [0023] Another objective of the present invention is that the Kalman filter processes the GPS phase measurements as well as the GPS pseudorange and delta range from both reference and rover sites, so as to improve the accuracy of the integrated positioning solution. [0024] Another objective of the present invention is that the Kalman filter is implemented in real time to optimally blend the GPS raw data and the INS solution and to estimate the navigation solution. [0025] Another further objective of the present invention is that a robust Kalman filter is implemented in real time to eliminate the possible instability of the integration solution. [0026] Another objective of the present invention is that a low accuracy inertial sensor is used to achieve a high accuracy integration solution by the aid of the global positioning system measurement. [0027] Another objective of the present invention is to provide a real-time integrated vehicle positioning method, which can substantially solve the problem of instability present in many existing systems where a Kalman filter is used to perform optimal estimation. [0028] Another objective of the present invention is to provide a real-time integrated vehicle positioning method, which supports high precision navigation in general aviation and space applications. It can also be used for ground motion vehicles tracking and navigation applications. [0029] Another objective of the present invention is to provide a real-time integrated vehicle positioning method, which uses the GPS raw phase measurements to update the inertial navigation system and aids the GPS tracking loop by the accurate output of the inertial navigation system so as to satisfy the requirements of, such as, low cost, high accuracy, continuous output, high degree of resistance to jamming, and high dynamics, and to overcome the disadvantages of the existing techniques. [0030]FIG. 1 is a block diagram illustrating an improved fully-coupled vehicle positioning method and system with differential GPS according to a preferred embodiment of the present invention, in which the global positioning system measurement and the inertial measurement are blended in a central navigation processor. [0031]FIG. 2 is a block diagram of the central integrated navigation processing, including the global positioning system and inertial sensors, according to the above preferred embodiment of the present invention. [0032]FIG. 3 is a flow diagram of the new process for on-the-fly ambiguity resolution technique of the present invention. [0033]FIG. 4 is a flow diagram of intermediate ambiguity search strategy (IASS) according to the new process for on-the-fly ambiguity resolution technique of the present invention. [0034]FIG. 5 is a block diagram of the procedure for forming the estimator bank according to the new process for on-the-fly ambiguity resolution technique of the present invention. [0035]FIG. 6 is a complete form of the estimator bank according to the new process for on-the-fly ambiguity resolution technique of the present invention. [0036]FIG. 7 is a block diagram of the inertial navigation system processing, which receives the navigation state corrections from a Kalman filter according to the above preferred embodiment of the present invention. [0037]FIG. 8 is a block diagram of the robust Kalman filter implementation according to the above preferred embodiment of the present invention. [0038]FIG. 9 is a block diagram of the master-slave relative positioning process according to the above preferred embodiment of the present invention. [0039]FIG. 10 is a block diagram of the GPS processor according to the above preferred embodiment of the present invention. [0040]FIG. 11 is a block diagram of the preferred attitude determination according to the above preferred embodiment of the present invention. [0041]FIG. 12 is a block diagram of the improved navigation application system according to the above preferred embodiment of the present invention. [0042] The improved fully-coupled GPS/IMU vehicle positioning system with differential GPS of the present invention, as shown in FIG. 1, comprises an IMU (inertial measurement unit) [0043] Referring to FIG. 1 and FIG. 2, the improved fully-coupled global positioning system/inertial measurement unit (GPS/IMU) vehicle positioning process with differential GPS of the present invention comprises the following steps. [0044] a) Receive GPS rover measurements (including pseudorange, carrier phase, and Doppler shift) from the GPS processor [0045] b) Combine the output of the INS (inertial navigation system) processor [0046] c) Feed back the output of the Kalman filter [0047] d) Inject the corrected velocity and acceleration data from the INS processor [0048] e) Inject the outputs of the GPS processor [0049] f) Output carrier phase integer ambiguities as the ambiguities are fixed from the on-the-fly ambiguity resolution module [0050] g) Output navigation data from the INS processor [0051] The master-slave positioning system comprises of a master system and a slave system, as shown in FIG. 9. The master system comprises a GPS antenna [0052] The master-slave relative positioning process comprises steps of: [0053] (a) Receive GPS signals by a GPS antenna [0054] (b) Shift the carrier frequency of the received GPS signals by the frequency mixer [0055] (c) Broadcast the carrier frequency shifted GPS signals by a data link [0056] (d) Receive GPS measurements (including pseudorange, carrier phase, and Doppler shift) from the GPS processor [0057] (e) Combine the output of the INS (inertial navigation system) processor [0058] (f) Feed back the output of the Kalman filter [0059] (g) Inject the corrected velocity and acceleration data from the INS processor [0060] (h) Inject the outputs of the GPS processor [0061] (i) Output carrier phase integer ambiguities as the ambiguities are fixed from the on-the-fly ambiguity resolution module [0062] (j) Output navigation data from the INS processor [0063] The master-slave navigation configuration performs autonomous navigation processing on a slave carrier, and determines the precise relative position with respect to a second one (master). These navigation functions are validated by executing kinematic differential GPS processing based on the observables extracted from the direct GPS signals and the GPS-like signals transmitted by a data link on the master carrier. [0064] A GPS antenna of the master system receives all the visible GPS satellite signals. These signals are transmitted through a data link after a carrier frequency shift procedure. All features of the GPS signal are maintained on the transmitting GPS-like signal, such as the C/A code modulation, P code modulation, navigation message modulation, etc. On the slave system, a corresponding data link is used to receive the GPS-like signals from the master spacecraft. Then an inverse carrier frequency shift is made on the slave carrier to retrieve the signal transmitted by the master system. The retrieved signals carry the master carrier's position and velocity information which can be solved on the slave carrier (as shown in FIG. 9). While the slave carrier receives the signals directly from the GPS satellites, the slave carrier can determine its position and velocity based on these direct measurements. Furthermore, carrier phase differential processing can be executed by the slave carrier to get the precise relative position between the master and slave carriers. [0065] Assuming the GPS signals received by the master carrier are transmitted without delay, the GPS-like signal received by the slave carrier can be represented as: ρ [0066] where, ρ [0067] In the above equation, ρ [0068] The position of the slave carrier can be determined from the direct GPS signals in the conventional manner. The relative position can be solved precisely from the differences in pseudorange and carrier phase between the direct GPS signal and those derived from the signals retransmitted by the master carrier. The equations are essentially identical to those of conventional kinematic positioning as the delay due to d [0069] Similarly, the velocity of the master carrier and the relative velocity between the master carrier and the slave carrier can be determined through the Doppler shift measurements. Time differentiation of the positioning equation gives: {dot over (ρ)} [0070] where: {dot over (ρ)} [0071] The mutual relative positioning (MRP) system can be obtained by combining the master and slave subsystems together and installing them on each carrier. [0072] The GPS receiver derives the range from the master carrier to the satellite plus the distance between the slave and master carriers through the determination of the GPS signal propagation delay. In a GPS receiver, the code delay-lock loop (DLL) is used to capture the GPS signal and measure the time shift. Two clocks are involved. One is the satellite clock tagging the signal emission time. The other is the receiver clock which records the signal reception time. An atomic time system, referred to as the GPS time, is applied to provide time reference for the whole system. The time shift measured by the code DLL is presented by Δ [0073] where t [0074] The bias δ [0075] In the above equation, ρ [0076] In the navigation processing, the second term of the most-right hand side of the above equation is compensated by the following procedure. First, we can get an approximate Δt by calculating the satellite position at epoch t [0077] In order to accurately calculate the pseudorange measurement, all error sources should be considered to correct the term ρ(t [0078] where, ρ [0079] Similarly, the carrier phase model for master-slave relative positioning is given by
[0080] where, N [0081] For satellite k, the carrier phase single difference (SD) results from the difference between the two equations in the above equation, as follows
[0082] where N [0083] The carrier phase double difference (DD) is a function of the carrier phase single difference:
[0084] where N [0085] The purpose of the corrected INS velocity-acceleration information aided GPS PLL loop is to estimate the carrier phase of the intermediate frequency signal θ θ [0086] The problem now becomes to estimate the parameters of the above equation. The velocity-acceleration information, which describes the flight vehicle dynamic characteristics, is translated into the line-of-sight (LOS) velocity-acceleration information. Therefore, the estimate of the carrier phase of the intermediate frequency signal can be formulated by LOS velocity-acceleration values as follows: {circumflex over (θ)}(t)= [0087] where (b [0088] V [0089] The code tracking loop of the GPS processor [0090] The central navigation processor [0091] The GPS processor [0092] The central navigation processor [0093] The INS processor [0094] The GPS processor [0095] The new satellites/cycle slips detection module [0096] The on-the-fly ambiguity resolution module [0097] The Kalman filter [0098] It is well known that the receiver measurement noise for the L [0099] For GPS measurements, the double difference equations for L [0100] where (•) [0101] is the double difference residual of the ionospheric effect for L [0102] respectively, and the corresponding integer ambiguities are
[0103] respectively. Therefor, the frequencies for the wide lane and narrow lane ambiguities are equal to ƒ [0104] and the approximated double difference narrow lane ambiguity (real number) is given by [0105] where Ñ [0106] denotes the ionospheric signal observation,
[0107] and λ [0108] respectively. [0109] The advantage of the IMU aiding phase ambiguity resolution and cycle slip detection is that the precision vehicle coordinates and velocity from the corrected INS solution are available to aid in determining the original ambiguities and the search volume. Additionally, the INS aiding signal tracking enhances the receiver's capability to hold the global positioning system satellite signal, thus the probability of signal loss or cycle slip is reduced. [0110] Referring to FIG. 2, the on-the-fly ambiguity resolution module [0111]FIGS. 3, 4, [0112] The on-the-fly ambiguity resolution module [0113] (a) initiating an on-the-fly ambiguity resolution module as the new satellites/cycle slips detection module is on, i.e., the new satellites or cycle slips occur; [0114] (b) fixing integer ambiguities to estimate a more accurate vehicle navigation solution, [0115] (c) sending the selected integer ambiguities from the on-the-fly ambiguity resolution module to the Kalman filter [0116] The above step (b) further comprises: [0117] (b.1) using intermediate ambiguity search strategy (IASS) and estimator bank to set up ambiguity set and determine the ambiguity integer; [0118] (b.2) validating and confirming the ambiguity integer. [0119] Basically, IASS comprises the “simplified” least-squares method and the extrawidelaning technique. Before using the least-squares method to search the ambiguities, the observable common satellites between two antennas (reference and rover) are divided into two groups: [0120] the primary satellites and the secondary satellites. Since the double difference equations are used, the satellite with the highest elevation is defined as the reference satellite. The primary satellites include the next four higher elevation satellites, i.e., there are four independent double difference equations. [0121] The rest of the observable satellites are categorized into the secondary satellites. [0122] As shown in FIG. 4, the IASS process comprises of a primary double difference wide lane ambiguity resolution module [0123] The first step of the IASS is to resolve the primary double difference wide lane ambiguities in the primary double difference wide lane ambiguity resolution module [0124] After the estimation of the primary double difference wide lane ambiguities, the estimated primary double difference wide lane ambiguities and the corresponding cofactor matrix are sent to the ambiguity domain determination module [0125] where {circumflex over (x)} [0126] The fixed primary double difference wide lane ambiguities are sent to the position calculation module [0127] Substituting the resolved double difference wide lane ambiguities into Equation (2), the approximated double difference narrow lane ambiguities (real numbers) are calculated. The extrawidelaning technique states that if the wide lane ambiguity is even (or odd), the corresponding narrow ambiguity is even (or odd), and vice versa. Using the extrawidelaning technique, the narrow lane ambiguities can be resolved in the extrawidelaning technique module [0128] and
[0129] respectively. [0130] Returning to FIG. 3, when the current ambiguity set from the IASS is different from the one(s) from the previous epoch(s), the current ambiguity set becomes a new member of an estimator bank [0131] 1. Search the integer ambiguity set at the first epoch of the search window by using the IASS. The integer ambiguity set becomes a member of the estimator bank [0132] where P [0133] and the first term of the product can be expressed as
[0134] which is assumed and defined as a Gaussian distribution. Equation (4) states the accumulative property of P [0135] Of course, the value of the only weight (D=1 in Equation (3)) in the weight bank [0136] 2. Search the ambiguity set by using the IASS at the second epoch of the search window. Two situations may occur: [0137] 2-1. When the integer ambiguity set is the same as the one of the previous epoch (epoch one), the number of the Kalman filters in the estimator bank [0138] 2-2. When the integer ambiguity set is different from the one of the previous epoch (epoch one), the current ambiguity set becomes a new member of the estimator bank [0139] 3. Follow the same procedure as described in step [0140] Referring to FIG. 3, after the period of the search window, still, the phase measurements (reference and rover) are input into the complete estimator bank P [0141] where C denotes a very large uncertainty to make sure that the ambiguity set is robust enough. After the criterion is met, the estimator bank [0142] Referring to FIG. 7, the INS processor [0143] The IMU I/O interface [0144] In addition to the corrected body angular rates from the IMU error compensation module [0145] where q [0146] where Δθ is the rotation angle and α,β, and γ are the angles between the axis of rotation and the axes of a coordinate system. For instance, they are the angles with respect to the roll, pitch, and yaw axes. Also, the quaternions satisfy the condition [0147] Ω [0148] Ω [0149] If the navigation frame is the local level North, East, and Down (NED) navigation frame, then
[0150] where ω [0151] The coordinate transformation module [0152] The attitude position velocity computation module [0153] where a and V are the acceleration and velocity of the vehicle relative to the Earth in the navigation frame, ω [0154] Because the accelerometers do not distinguish between vehicle acceleration and the mass attraction gravity, the specific vector, f, sensed by the accelerometers is: [0155] where g(r) is a combination of the earth's gravity and the centrifugal force at the vehicle location. Thus, [0156] where,
[0157] The vehicle velocity is updated by the following: [0158] where C [0159] For the WGS- [0160] where m=ω [0161] The differential equations for the position update of the geodetic latitude, L, longitude, λ, and height, h, are given by:
[0162] where R [0163] After the computation of the position and velocity, the position and velocity errors calculated by the Kalman filter [0164] The corrected inertial solution obtained from the attitude position velocity computation module [0165] It is well known that the Kalman filter produces optimal estimates with well defined statistical properties. The estimates are unbiased and they have minimum variance within the class of linear unbiased estimates. The quality of the estimates is however only guaranteed as long as the assumptions underlying the mathematical model hold. Any misspecification in the model may invalidate the results of filtering and thus also any conclusion based on them. [0166] In the improved real-time fully-coupled GPS/IMU positioning process and system with differential GPS, an alternative for a Kalman filter for position and attitude derivation is a robust Kalman filter. This robust Kalman filter is stable enough to operate in more than one dynamical environment. If the dynamics change drastically, or if a sensor failure occurs, for example, a GPS satellite signal failure or an inertial sensor signal failure, the filter must detect, rectify and isolate the failure situation. [0167] A robust filter has the characteristic that it provides near-optimum performance over a large class of process and measurement models. The pure Kalman filter is not robust since it is optimal for only one particular process and measurement model. If the filter is not correct, the filter covariance may report accuracy which is different from what can actually be achieved. The purpose of filter integrity is to ensure that the predicted performance from the error covariance is close to the actual estimation error statistics. In addition, filter divergence is usually caused by a changing process, measurement model, or a sensor failure. [0168] This present invention uses a residual monitoring method to obtain a robust Kalman filter which is used to blend the global positioning system raw data and the inertial sensor measurements. When the proper redundancy is available, residual monitoring schemes can efficiently detect hard and soft failures and filter divergence. One benefit of the residual monitoring approach is that when the filter model is correct, the statistical distribution of the residual sequence is known. Thus, it is easy to generate a measurement editing and divergence detection scheme using a test-of-distribution on the measurement residuals. The same statistics can be used to assess the filter tuning and adjust the size of the covariance when divergence is detected. FIG. 8 gives the implementation of this robust Kalman filter including a residual monitoring function. [0169] As shown in FIG. 8, a GPS error compensation module [0170] A preprocessing module [0171] The state vector prediction module [0172] The computing measurement residue module [0173] The residue monitor module [0174] The covariance propagation module [0175] The computing optimal gain module [0176] The updating state vector module [0177] In view of the above, the present invention can provide a real-time fully-coupled vehicle positioning process and system with differential GPS to substantially solve the problems encountered in global positioning system-only and inertial navigation system-only, such as loss of global positioning satellite signal, sensitivity to jamming and spoofing, and the inertial solution's drift over time. Therefore, the following features and advantages can thus be achieved: [0178] (1) The velocity and acceleration from an inertial navigation processor are used to aid the code and carrier phase tracking of the global positioning system satellite signals, so as to enhance the performance of the global positioning and inertial integration system, even in heavy jamming and high dynamic environments. [0179] (2) The velocity and acceleration from an inertial navigation processor are corrected by a Kalman filter and used to aid the code and carrier phase tracking of the global positioning system satellite signals, so as to enhance the performance of the global positioning and inertial integration system, even in heavy jamming and high dynamic environments. [0180] (3) To accurately determine the receiver position and velocity at the centimeter level, the GPS phase measurements will be used and the differential GPS will be employed. In this invention, a new process (OTF (on-the-fly) technique) is proposed to resolve the integer ambiguities on the fly and estimate the receiver position in real time. The results of GPS estimates will increase the accuracy of the inertial navigation system and therefore enhance the capability of the GPS tracking loop. [0181] (4) To perform the differential GPS, the data link [0182] (5) The self-contained INS complements the GPS as the GPS receiver suffers the loss of the GPS signals. Once the GPS receiver regains the signals and then estimates the receiver position and velocity, the output (position and velocity) of the GPS receiver is used to correct the position and velocity of the INS that has drifted. [0183] (6) The inertial navigation system aids the satellite signal integer ambiguity resolution of the global positioning system by providing more accurate position information. [0184] (7) The integrated navigation solution of the global positioning system and the inertial measurement unit aids the satellite signal integer ambiguity resolution of the global positioning system by providing more accurate position information. [0185] (8) The satellite signal carrier phase measurements (reference and rover) are used in the Kalman filter, as well as the pseudorange and delta range of the global positioning system, so as to improve the accuracy of the integration positioning solution. [0186] (9) The Kalman filter is implemented in real time to optimally blend the global positioning system raw data and the inertial navigation solution and to estimate the navigation solution. [0187] (10) The robust Kalman filter is implemented in real time to eliminate the possible instability of the integration solution. [0188] (11) Low accuracy inertial sensor is used for achieving a high accuracy integration solution due to the aiding of the global positioning system measurements. [0189] The present invention can be used for Wide Area and Local Area Augmentation Systems where precision corrections are required. The formal requirements established by the FAA for the WAAS (Wide Area Augmentation System) stipulate that aircraft real time position should be determined to an accuracy of 7.6 meters in both the vertical and horizontal component planes with a 95% probability (two sigma). This assumes a conventional single-frequency user applying WAAS-supplied corrections to broadcast GPS orbits and clocks and to the ionopheric delay model. [0190] The wide-area master station (WMS) tracks and processes the GPS data to derive a vector correction for each GPS satellite. The vector correction will include GPS ephemeris errors, satellite clock errors, and ionospheric delay estimates. The FAA distinguishes two kinds of corrections: a slow correction and a fast correction. The slow correction contains the slowly varying errors: the ephemeris error and long-term clock errors. Due to its slowly varying nature, this error need be transmitted no more than every 120 seconds. On the other hand, the satellite clock error is quickly varying due to selective availability (SA). A given satellite should not be unreported for more than six seconds. The Message Type [0191] The present invention can be applied to autonomous navigation for Reusable Launch Vehicles (RLVs), which is used to transport payloads and humans between the earth and space successfully and safely. The autonomous multi-antenna GPS/DGPS/MEMS IMU/WAAS/Radar Altimeter integrated system provides precise, reliable indispensable navigation information for all of the RLV's flight phases. The real time automatic reconfiguration capability is realized by a flexible Federated Kalman filtering mechanization and an expert mission planing system. The configured GPS/INS integration provides a navigation solution during the RLV's ascent and descent phases. The configured GPS Interferometer/INS integration provides a navigation solution for the RLV's orbital operations, while the configured GPS/DGPS/WAAS/INS/Radar Altimeter integration provides a precise approach and landing capability for the RLV. The GPS, INS, Radar Altimeter, and WAAS sensors assume complementary roles so that optimal system performance is achieved for every flight phase with the characteristics of high precision, high integrity and seamless navigation. An intelligent neural network is applied to perform multi-sensor failure detection and isolation, and redundancy management. [0192] The boost stage is from take-off until the main engine cut-off. Since the launch vehicle may meet large disturbances, such as winds, structure disturbances, and thrust direction errors, launch vehicle attitude stabilization is of critical importance during this stage. Another problem is a potential failure of a GPS-only system due to high launch vehicle dynamics during the boost stage. The Fully-Coupled GPS/INS integrated system using pseudorange and delta range measurements provides robust navigation performance during the boost stage. The velocity and acceleration (V-A) information derived from the INS is injected into the GPS signal tracking loops to improve the dynamic performance and interference tolerance of the GPS receiver. [0193] Within the fully-coupled GPS/INS integration architecture, the MEMS IMU sensor and the GPS receiver complement each other at an advanced level. The advantages of the Fully-Coupled integration mode include: 1) hardware-level redundancy, 2) low-cost IMU sensors, 3) enhanced anti-interference, 4) dramatically extended dynamic range, 5) shortened time-to-first-fix (TTFF) and signal reacquisition time, and 6) excellent navigation accuracy. [0194] A low-cost multi-antenna GPS and MEMS IMU integration with interferometric processing executes navigation with an attitude estimation capability for the RLV's orbital operations. This integration approach eliminates the need for a star tracker for attitude estimation. GPS attitude determination using carrier phase measurements can be made with 3 antennas at the ends of two baselines of known length. Before attitude information can be computed, the initial phase ambiguities must be solved. [0195] A GPS/DGPS/WAAS/MEMS IMU/Radar Altimeter integration mode provides precise approach and landing capability for the RLV. The absolute positioning accuracy achievable from GPS is dependent upon the accuracy of the GPS measurements which are affected by many errors. The major error sources include ephemeris error, clock error including Selective Availability (SA), atmospheric effects and multipath. SA is the intentional degradation of the system imposed by the DoD through satellite clock dithering and is the largest error source using GPS. The GPS is augmented by the WAAS/DGPS data and the altitude measurements coming from a radar altimeter. [0196] During the descent stage, the fully-coupled GPS/MEMS IMU provides the RLV's position and attitude information under the high dynamic environment. Toward the end of the descent stage, the WAAS receiver starts to track the WAAS signal, and the DGPS receiver also starts to receive correction data from the ground-based reference station. In the DGPS/WAAS mode, the RLV navigation system is more reliable than DGPS-only or the WAAS-only system. A key consideration for DGPS/WAAS mode is to select a relevant ground reference station which provides the most applicable error corrections. A tradeoff analysis and decision should be made for the geometry. An approach that weighs the corrections based on the RLV-GPS-ground reference station geometry seems to be appropriate. [0197] Another typical application of present invention is in the aircraft industry, as shown in FIG. 12. The navigation sensor array comprises of an IMU [0198] A mission planner [0199] The preferred embodiment of the GPS processor [0200] (1) a RF unit [0201] (2) the correlation and tracking loops [0202] (3) the satellite and antenna selection module [0203] (4) the attitude determination processor [0204] (1) an integer ambiguity resolution module [0205] (2) a state update module [0206] (3) a state prediction module [0207] (4) the fuzzy logic inference module [0208] (5) a vehicle attitude propagation module [0209] (6) the phase difference prediction module [0210] (1) a fault detection and isolation (FDI) module [0211] (2) a filter gain computation [0212] The acceleration and velocity data from the central navigation and control processor [0213] The attitude data from the central navigation and control processor Referenced by
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