Showing papers on "GPS/INS published in 1992"

GPS tracking system

Abstract: A low cost tracking system employing satellites of the global positioning system (GPS) is suitable for applications involving radiosondes, sonobuoys, and other objects. The tracking system includes a sensor mounted on each object which digitally samples the GPS satellite signals and records them in a data buffer. The digital samples are then transmitted, at a rate lower than that at which the GPS satellite signals were sampled, over a data telemetry link, interleaved with other telemetry data from the object. The GPS data is processed in a data processing workstation where the position and velocity of the sensor, at the time the data was sampled, is computed. The data buffer in the sensor is periodically refreshed, and the workstation periodically computes the new position and velocity of the sensor. Differential corrections are also provided at the workstation to aid in signal acquisition and to increase the precision of the position fix.

Method and apparatus for improving the accuracy of position estimates in a satellite based navigation system

Abstract: An apparatus and method are disclosed for determining the position of a vehicle at or near the surface of the Earth using navigation signals from a satellite-based navigation system. Precise position estimates are achieved by reducing effective receiver noise. A plurality of receiver means each compute a pseudorange for each satellite. A signal optimizer then uses the pseudoranges from each receiver means to compute an optimal pseudorange for each satellite. A processor means computes a vehicle position from the optimal pseudoranges.

On-board navigation apparatus

Abstract: An on-board navigation apparatus in which coordinate data of formerly set designations are held in a memory, the destination coordinate data stored in the memory are read out at the time of destination setting, and one piece of destination coordinate data is selected from the read designation coordinate data in accordance with an operation input to thereby set a destination according to the selected one piece of destination coordinate data. The arrangement can allow a user to easily set the same destination as a formerly set destination only by a simple selecting operation.

Integrated enroute and approach guidance system for aircraft

Abstract: Data from long range aids such as the global positioning system (GPS) and an inertial navigation system (INS) and short range aids such as a microwave londing system (MLS) are used to smoothly and automatically transition an aircraft from the long range aids to the short range aids. During cruise a Kalman filter (70) combines data from the global positioning system and the inertial navigation system to provide accurate enroute information. When the aircraft arrives in the vicinity of the airport and begins to acquire (70) data from the microwave landing system, the Kalman filter (70) is calibrated with the MLS data to permit precision landing with GPS/INS data alone in case the MLS system subsequently fails. In addition, navigation information begins to be derived from a weighted sum of the GPS/INS and MLS data, the weighting being determined by distance from the airport. In a first region farthest from the airport, the GPS/INS data is given a 1.0 weighting factor; and in a second region nearest the airport, the MLS data is given a 1.0 weighting factor. In a third region intermediate the first and second regions, the GPS/INS data and MLS data are proportionately and complementarily weighted as a function of the distance from the airport. If the MLS system fails, the weighting system (76, 78, 80) is disabled (84) and navigation data is again derived from the GPS/INS combination. In addition, the data from both systems are monitored (74), and a cockpit alarm is sounded if the data diverged beyond a specific amount. The Kalman filter (70) supplies a weighting system comprising two weighting devices (76, 78), one (76) of which weights the MLS data, and the other (78) the GPS/INS combination data. The outputs of these devices (76, 78) are summed (80) and supplied to one terminal of a switch (84) for selecting between the weighting system (76, 78, 80) and the Kalman filter (70).

Method and apparatus for GPS positioning, filtering and integration

Abstract: A method and apparatus for processing data from a positioning or navigation system, such as global positioning system (GPS) navigation system data. In preferred embodiments, the invention is a method and apparatus for estimating position based on GPS measurements and optionally, measurements from other types of navigation sensors combined with GPS navigation system data. The apparatus of the invention includes a point position estimator (which preferably approximates a maximum likelihood estimator) for processing data from a positioning or navigation system to generate point position estimate data, and a smoothing filter for smoothing the point position estimate data. The smoothing filter can be a linear Kalman filter.

Kalman filter mechanization for INS airstart

Abstract: A strapdown mechanization and associated Kalman filter are developed to provide both ground align and airstart capabilities for inertial navigation systems (INSs) using Doppler velocity and position fixes, while not requiring an initial heading estimate. Position update during coarse mode is possible by defining sine and cosine of wander angle as filter states and modeling the position error in geographic frame while integrating velocity in the wander frame. INS Global Positioning System (GPS) differential position due to GPS antenna moment arm can aid heading convergence during hover turns in helicopter applications. Azimuth error state in the fine mode of the filter is defined as wander angle error to provide continuous estimation of navigational states, as well as inertial/aiding sensor errors, across the coarse-to-fine mode transition. Though motivated by a tactical helicopter application, the design can be applied to other vehicles. Advantages over conventional systems in addition to the airstart capability include robustness and versatility in handling many different operational conditions. >

Architectures and GPS/INS integration: impact on mission accomplishment

Abstract: It is pointed out that the GPS (Global Positioning System) signal contains information which, when properly combined with information from INS (inertial navigation system) and other sensors, provides exceptionally high-accuracy position, velocity, attitude, and time information. These ten elements of information (three each in position, velocity, and attitude, and one in time) are common, in various combinations, to most of the avionics functions. When viewed from a system perspective, this high-precision information can be thought of as the integration basis, or a reference set, which offers opportunities for reconfiguration of the offensive, defensive, communication, navigation, and other sensors. Various integration architectures for fusion of these sensors can inherently enhance, enable, or severely limit these potential mission capabilities. The choice of integration architecture can directly and profoundly affect performance, cost of integration, cost of ownership, and exploitation of much greater mission capability. This is illustrated with the GPS/INS integration example. >

Achieving modularity with tightly-coupled GPS/INS

Abstract: Addresses a problem of GPS/INS (Global Positioning System/inertial navigation system) integration wherein the conflicting goals of modularity and tight coupling are both sought. The advantages to be gained from modularity are described. The performance advantages of tight coupling are reviewed, and it is explained why it is so difficult to achieve both modularity and tight coupling in the same architecture. A candidate is then presented and evaluated relative to the stated goals. Key aspects of its operation and timing structure are described, including difficulties the architecture overcomes such as discontinuous INS aiding of receiver tracking loops as the system transitions through various modes. Computational requirements and throughput or the integration processor are discussed. The problem of testing individual components is considered where those components are normally expected to operate only when fully integrated. The author deals primarily with low-cost applications such as remotely piloted vehicles and tactical munitions where tight coupling and modularity are especially valuable. >

Differential GPS/inertial navigation approach/landing flight test results

Abstract: In November 1990, a differential GPS/inertial flight test was conducted to acquire a system performance database and demonstrate automatic landing using an integrated differential GPS/INS with barometric and radar altimeters. Flight test results obtained from postflight data analysis are presented. These results include characteristics of DGPS/inertial error, using a laser tracker as a reference. In addition, data are provided on the magnitude of the differential correlations and vertical channel performance with and without radar altimeter augmentation. Flight test results show one sigma DGPS/inertial horizontal errors of 9 ft and one sigma DGPS inertial vertical errors of 15 ft. Without selective availability effects, the differential corrections are less than 10 ft and are dominated by receiver unique errors over the time period of an approach. Therefore, the one sigma performance of the autonomous GPS (8-ft horizontal and 20-ft vertical) is very similar to the DGPS/inertial performance. Postprocessed results also demonstrate significant improvements in vertical channel performance when GPS/inertial is aided with radar altimeter along with a low-resolution terrain map. >

The GPS filtering problem

Abstract: The authors explore the possibility of both improved navigation accuracy and computational simplification by using nonlinear filters based on direct solutions to the GPS (Global Positioning System) equations. After reviewing results concerning the existence of sufficient statistics for the nonlinear GPS filtering problem, they introduce the notion of a two-stage estimator in which a direct solution is combined with a time-series smoothing algorithm, such as a constant-gain Kalman filter. This method provides a means for decoupling, in a sense, the spatial and temporal aspects of the GPS filtering problem. Experiments using real data suggest that the method has advantages over the extended filter, in terms of both computational burden and accuracy. >

Extracting gravity vectors from the integration of global positioning system and inertial navigation system data

Abstract: Consulting records of a past flight, 1-Hz trajectory, and (roll, pitch, yaw) motion profiles were adopted for a simulated 7.5-hour balloon flight over New Mexico. The assumed pay load of the balloon was an inertial ring laser gyro strap-down system and a three-antenna, minimum 12-channel Global Positioning System (GPS) receiver. The adopted trajectory was assumed to be GPS determined and interferometric in nature reflecting a second, ground-fixed receiver. Owing to the strap-down Inertial Navigation System (INS), a corresponding 1-Hz profile of accelerometer specific force outputs (total (GPS sensitive) inertially referenced accelerations less gravitational accelerations present) was also required to conduct a covariance error analysis. The gravitational accelerations were computed by a spherical harmonic expansion through n = m = 360. A 36-state, 40-noise process, open-loop Kaiman filter integrating GPS and INS data was constructed. A full constellation of 18 GPS satellites was simulated. Related to the two receivers, external GPS updating observation types used were pseudorange differences, single differenced carrier phases, and single differenced phase rates. A fourth type of updating observation used was the error in the INS's estimation of phase differences between GPS antenna pairs (an equilateral triangle balloon-borne configuration was assumed with 1-m baselines). The filter cycle and update frequency were both set at 1 Hz. A postmission covariance analysis based Kaiman filter mechanization was executed, taking into account the major initial INS and GPS error sources. Hybrid alignment errors were reduced to 0.4 arc sec about the north and east axes (leveling errors) and to 0.7 arc sec about the down axis (heading error). Hybrid positioning and velocity errors were held to the 1–2 cm and 1–2 mm\s range, respectively, and were highly correlated to the corresponding position dilution of precision satellite geometry. The benign motions of the balloon allowed for a smooth time behavior of the states. Assuming the accelerometer errors are essentially bias states that can be calibrated in flight based on very accurate GPS observations, the leveling uncertainties are the main error sources in obtaining the three components of the balloon- borne gravity vector. The alignment errors mentioned above contribute a 2.5-mGal error to the computed horizontal components and a 0.05-mGal error to the vertical. Limiting the misalignment contributions to the deflection error budgets to 1 mGal will require leveling errors of 0.2 arc sec. Gravity components obtained from a hybrid system can be compared to upward continued (from extensive ground gravity data) components to validate continuation models and to establish instrumental proof of concept. A successful balloon experiment would be a strong argument for a “high-low” satellite tracking mission involving the GPS constellation and a dedicated, polar orbit, low satellite (altitude e (160 km, 300 km)) possessing the same payload. At such an altitude the leveling concerns are less pronounced. Such a spaceborne mission would allow for a downward continued gravity mapping over all heretofore inaccessible Earth surface regions.

Integrated GPS‐INS for High‐Accuracy Road Positioning

Abstract: The use of a differential global positioning system (GPS) for high‐accuracy road positioning has shown some limitations in terms of outages due to satellite masking that cause poor geometry and loss of carrier phase lock. By integrating a GPS receiver to an inertial navigation system (INS), i.e., GPS/INS, these limitations can be overcome to a large degree. The concept of GPS/INS for positioning is discussed with the emphasis on accuracies at the centimeter level, and an effective cycle‐slip detection and correction strategy using INS is also presented. GPS/INS data collected on a highway near Calgary is used to assess the feasibility of the system in terms of accuracy and reliability. GPS outages are simulated to determine their effect on the accuracy and cycle‐slip detection/correction capabilities. Results show that accuracies of 5 cm are achievable with this system; however, a degradation in performance can be expected when GPS outages occur. A discussion of improved modeling techniques as well as the...

Balloon gravimetry using GPS and INS

Abstract: The US Air Force has a project to measure the horizontal components of gravity at altitude using balloon-borne instrumentation consisting of a Global Positioning System (GPS) receiver and a strapdown inertial navigation system (INS). GPS data are to be used primarily to determine the total inertial acceleration of the balloon, while the INS accelerometers sense all nongravitational accelerations. A covariance analysis based on the Kalman filter shows that conventional gravity estimation from GPS-aided INS data is possible only if external attitude updates are available. An alternative technique is explored that attempts to estimate at least part of the gravitational spectrum without modeling the gravity disturbance as a state variable or relying on external attitude updates while, at the same time, admitting uncorrected (long-wavelength) attitude errors. Simulations based on a model for typical balloon motion are used to discuss this possibility. >

Global positioning system based method for high-accuracy and high-confidence level absolute and relative position and velocity determinations

Abstract: An embodiment of the present invention is a shipboard GPS positioning system (12) having data links (28, 30, 32, 34) to outlying tailbuoys (18, 20, 22, 24) requipped with respective GPS receivers On ship (12), an Intel 386-based microcomputer system collects data from various ship's equipment including the ship's GPS receiver and data from the several tailbuoy units (18, 20, 22, 24) A computer-implemented process located in the microcomputer system controls the following processes Periodically each GPS receiver produces updated pseudo ranges (PRs) and these are time-tagged The time-tagged PRs for the ship (12) are aligned according to their time tags with their counterpart PRs from the tailbuoys (18, 20, 22, 24) The raw PRs are then passed through a Kalman mathematical filter to produce filtered pseudo ranges A position solution is then attempted for each GPS receiver using the The filters provide statistical data that is used to rate the quality of each in a weighted least squares solution process Special measures are included to provide quality control/quality assurance, including the user of error ellipses on the display to present a graphic indication of expected solution accuracy, and the use of real time fault detection, isolation, and correction algorithms when redundant satellite information is available

Gravity vector estimation from integrated GPS/strapdown IMU data

Abstract: The Air Force is planning to launch a helium-filled balloon possessing a payload of an inertial ring laser gyro strapdown system and a 3-antenna, minimum 12-channel, dual-frequency GPS receiver. A ground-fixed, single-antenna GPS receiver will be centrally placed beneath the balloon’s trajectory, allowing for differential mode GPS tracking. The inertial measurement unit’s (IMU’s) accelerometers will measure the total kinematic accelerations in the body frame. These will be transformed to local (n,e,d) navigation frame components using the integrated attitude of the balloon obtained from the IMU’s gyroscopes. GPS-determined accelerations will reflect total inertial accelerations, which can easily be expressed in the navigation frame. Differencing the two sets will thus yield the desired gravitational accelerations. An elaborate covariance study was conducted involving a 36-state, 43-noise-process open-loop Kalman filter. Using conservative initial IMU and GPS error source values, the covariance study suggests that estimation of all three components of balloon-borne gravity vectors to an accuracy of 5 mGa1 (1 mGa1 = 10~5 m/s2) is feasible using “within reach” electronic and engineering capabilities. Leveling uncertainties are the main contributors to the overall gravity vector estimation error budget.

GPS navigation system with local speed direction sensing and PDOP accuracy evaluation

Abstract: GPS present position and/or speed sensor distance data corrective coefficient are updated depending on accuracy evaluation, minimum speed, system independency, uninterrupted data flow from satellite and statistical error: A GPS (Global Positioning System) navigation system for a vehicle determines the present position from GPS radio waves (1) and from direction (6) and speed (7) sensor system (2). A data processor (9) selects updating of present position from the GPS device (1) and/or corrects a distance corrective coefficient for the speed sensor (7) data if five conditions are fulfilled: a) if a PODP accuracy evaluation value of GPS (1) data is equal or smaller b) if GPS speed data are equal or greater c) if the difference between GPS and sensor speed data is equal or smaller than respective predetermined values and d) if GPS data were obtained within three consecutive time intervals of predetermined lenght and e) if these three consecutive GPS data are equal or smaller than a predetermined value.

Determining inertial errors from navigation-in-place data

Abstract: To validate the self-calibration performance of a missile system's inertial navigation system (INS), an inertial error estimator (IEE) and a calibration procedure were developed and tested. The IEE was developed using a least squares fit of navigation-in-place data to calculate the accelerometer biases, the gyroscope biases, and the axis misalignments. It was found that the estimation error due to highly correlated regression functions decreases with navigation time. A tradeoff study was performed to select the optimum navigation-in-place time allowed for determining the estimation of the inertial errors. Once the optimum navigation time was determined, an INS calibration procedure incorporating the IEE was developed. The calibration procedure was validated by applying corrections for the inertial errors to the INS, and then navigating in place to show a reduced position error profile. Test results showing the outcome of the INS calibration procedure are included. >

Maintaining high accuracy GPS positioning 'on the fly'

Abstract: A series of airborne GPS (Global Positioning System) tests using the Ashtech XII receivers was used to demonstrate the effect of cycle slips and erroneous data on the positioning accuracy. The overall achievable accuracy is in the order 20 cm; however, errors of several decimeters were detected in some instances. Traverse closures and residual analysis are used to assess the occurrence and magnitude of these errors. The addition of an inertial navigation system in the aircraft, which provides accurate relative positions, is also used to detect cycle slips and outliers in the GPS data. Strategies to achieve consistent accuracies are discussed, and results for different alternatives are presented. >

Detection of Spoofing, Jamming, or Failure of a Global Positioning System(GPS)

Abstract: : The Air Force has equipped its aircraft with avionic systems such as Global Positioning System (GPS) and Inertial Guidance Systems (INS) capable of providing accurate navigation solutions. The aircrews flying these aircraft require a system that can either survive the hostile environments encountered in combat or notify the aircrew that their performance has been significantly degraded. This research focuses on failure detection and isolation techniques using an extended Kalman filter and generalized likelihood ratios using matched filters. Analysis is conducted using a Kalman filter development package known as the Multimode Simulation for Optimal Filter Evaluation (MSOFE). Both a large order truth model for the navigation system (in which a full 24 satellite constellations is modeled) and a reduced-order Kalman filter are developed. Results suggest that failures within the GPS can be detected, isolated, and in some cases compensated through feedback.... Spoofing, Failure detection, Kalman filter, Jamming.

Low cost inertial measuring unit

Abstract: A concept being explored is the use of an inexpensive IMU (inertial measuring unit) with the GPS (Global Positioning System). An inexpensive IMU can meet many of the same functional requirements as the high-accuracy system, with the exception of long-term position. For example, some applications for instrumenting a reference frame are antenna stability, satellite acquisition (including GPS), platform stability, flight stabilization, and heading information. The acceleration data could still be integrated to find velocity and position; however, this position information would only be used for a period of minutes. For example, to provide continuous navigation during a GPS blockage due to buildings, wing blockage during a maneuver, etc. The author presents the capabilities of a unit of this type which was developed for the US Army. >

Speed-sensor-coefficient computing apparatus

Abstract: PURPOSE:To improve position measuring accuracy by computing the ratio between the straight distance between two GPS positions, which are measured with a GPS, and the straight running distance based on the output value of a speed sensor in the straight running on the straight distance, and obtaining the coefficient of the speed sensor. CONSTITUTION:An estimated-position computing means 104 computes the estimated position and the estimated bearing of a vehicle based on the outputs from a speed sensor 101 and an angular velocity sensor 102. A straight-running detecting means 105 detects the straight running of the vehicle based on the computed estimated position and estimated bearing. A speed-sensor-coefficient computing means 106 computes the speed sensor coefficient based on the ratio between the distance between two GPS position, which are measured with a GPS in straight running, and the distance based on the output value of the speed sensor 101, which is obtained during the period from the time when the first GPS position of two GPS positions is measured to the time when the second GPS position is measured. The speed sensor coefficient can be automatically corrected. Therefore, it is not necessary to correct the speed sensor coefficient directly by the user of the navigation apparatus in response to time change. The position measuring accuracy is also improved.

Compensation of selected availability using a GPS/INS extended Kalman filter

Abstract: The authors describe research to compensate for selected availability (SA) corruption by modifying the Kalman filter used in a typical Global Positioning System/inertial navigation system (GPS/INS) integration. Two modified filters are developed: one filter that models SA using individual error-states and another filter that compensates for SA by increasing the measurement noise. Neither of these filters totally eliminates the corruption due to SA, but both provide a significant increase in the accuracy of the navigation solution. >

System for determining the position of mobile objects

Abstract: In order to determine the position of mobile objects (5), in particular vehicles, the invention uses Global Positioning System (GPS) satellites (1, 2, 3, 4) which orbit in polar orbits, continuously changing their position with respect to a stationary point on the Earth's surface and transmitting GPS signals. These GPS signals make it possible for a GPS receiver installed in the mobile object to determine its instantaneous position and altitude coordinates. In addition to the mobile GPS receiver, the invention calls for at least one GPS receiver (6) which operates as a reference receiver and whose position coordinates are known precisely. Connected to the mobile GPS receiver is a processing unit which continuously determines error values from the signals received continuously by the reference GPS receiver and from the known position coordinates of its location. The continuously determined error values are conditioned and transmitted (7) by a transmitter. The transmitted conditioned error values are continuously received in the mobile object (5) and analysed in such a way that the results produced by the GPS receiver installed in the mobile object are corrected by an amount corresponding to the error values. A radio transmitter is used as the transmitter, the conditioned error values being inserted into the programme signal broadcast by this radio transmitter. Also fitted in the mobile object is a radio receiver which receives the programme signal with the conditioned error values inserted in it, and a series-connected device for separating the conditioned error values from the programme signal.

Long duration strapdown stellar-inertial navigation using satellite tracking

Abstract: A strapdown stellar-inertial navigation system (INS) is currently being developed which can be used to correct for the gyro drift in an INS throughout a long-duration mission. The star sensor provides updates of the inertial angle error derived from stellar observations. However, since the star sensor updates are unable to correct for INS drifts due to accelerometer errors, the stellar inertial navigation solution accuracy degrades with time. A study has been performed on the number of visible satellites that can be tracked by the stellar-inertial system. The authors discuss the results of this study and include simulation results showing the performance improvement possible using both star and satellite observations to bound the system error growth. The results have shown that the combination of measurements can maintain the system error growth to within 700 ft over a long-duration mission. >

Real Time Estimation Of GPS Measurement Error From Vehicle Position Data

Abstract: Real time estimation of the accuracy of the Global Positioning System ( GPS) measurement is a key factor in vehicle applications. While the possibility of selective availability exists, estimation of the GPS total error is of fundamental importance. The algorithm proposed in this paper estimates this error showing good characteristics, and providing a useful tool for the test of the performance during working of GPS in a land vehicle.

GPS-INS data integration for remote sensing

Abstract: Summary form only given, as follows. Flight tests have been performed for the purpose of assessing the use of post-mission integrated data from GPS (Global Positioning System) and INS (inertial navigation system) systems for remote sensing applications. While GPS data alone are sufficient for aircraft guidance, they do not currently satisfy the requirement for position and attitude information at high data rates, with typical sampling times between 5 and 40 ms. This information is needed for image processing applications where data discontinuities greater than a few centimeters in the positions must be avoided. The authors describe the test design which provided for multiple GPS stations on the ground from which the GPS receiver in the aircraft was monitored. The airborne GPS receiver was precisely synchronized with the stable platform INS. The accuracy of position, velocity, and attitude was analyzed using different approaches to post-mission data integration. These results were then used to georeference the imagery from the Multiple-Detector Electro-Optical Imaging Scanner and to compare it to independent ground truth. >

Review of Techniques for In-Flight Transfer Alignment

Abstract: : This report presents methods for transfer alignment (TA) of inertial navigation systems (INS) which have been published in the accessible literature during the last three decades. Kalman filtering techniques based on linearised dynamics dominated in the literature of the subject. Methods of TA can be classified as angular rates, velocity or position matching. In each case a number of assumptions are made to ensure the validity of proposed technique. The accuracy of filters depends on the particular implementation viz. allowable maneuvers and time of TA, microprocessor used, vibration environment, inclusion of wing flexure into the model, type of application under discussion (range of missile or time of flight), quality of output from inertial measurement units, etc. Authors briefly discuss methods of in-flight transfer alignment of INS taking into account these assumptions. Inertial navigation, Inertial measurement and units, Calibration, Guided weapons.

Evaluating the Velocity Accuracy of an Integrated GPSANS System: Flight Test Results

Abstract: Verifying the velocity accuracy of a GPS receiver or an integrated GPS/INS system in a dynamic environment is a difficult proposition when many of the commonly used reference systems have velocity uncertainities of the same order of magnitude or greater than the GPS system. The results of flight tests aboard an aircraft in which multiple reference systems simultaneously collected data to evaluate the accuracy of an integrated GPS/INS system are reported. Emphasis is placed on obtaining high accuracy estimates of the velocity error of the integrated system in order to verify that velocity accuracy is maintained during both linear and circular trajectories. Three different reference systems operating in parallel during flight tests are used to independently determine the position and velocity of an aircraft in flight. They are a transponder/interrogator ranging system, a laser tracker, and GPS carrier phase processing. Results obtained from these reference systems are compared against each other and against an integrated real time differential based GPS/INS system to arrive at a set of conclusions about the accuracy of the integrated system.

Missing relativity terms in gps

Abstract: The current GPS relativity corrections were derived by a committee of experts chosen by the Air Force. The original derivation assumed that the GPS reference frame is inertial, with its origin at the center of the earth. It also assumed that the user obtains inertial GPS coordinate time from the satellites. A new approach in relativity is taken to determine the relativity effects for both GPS satellites and GPS receivers. To the first-order approximation, there are no new relativity effects missing in the GPS algorithm if one converts all observations to the earth-centered inertial (ECI) frame and the GPS inertial coordinate time. However, there are some uncompensated effects required to make the complete transformation from the noninertial frame of a GPS receiver to the inertial frame of GPS. These effects and their magnitudes are examined in this paper.