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This invention is drawn to an autonomous navigation system using Global Positioning System (GPS) and magnetometers for low Earth orbit satellites. As a magnetometer is reliable and always provides information on spacecraft attitude, rate, and orbit, the magnetometer-GPS configuration solves GPS initialization problem, decreasing the convergence time for navigation estimate and improving the overall accuracy. Eventually the magnetometer-GPS configuration enables the system to avoid costly and inherently less reliable gyro for rate estimation. Being autonomous, this invention would provide for black-box spacecraft navigation, producing attitude, orbit, and rate estimates without any ground input with high accuracy and reliability.

This paper documents the testing and development of magnetometer-based gyroless attitude and rate estimation algorithms for the Far Ultraviolet Spectroscopic Explorer (FUSE). The results of two approaches are presented, one relies on a kinematic model for propagation, a method used in aircraft tracking, and the other is a pseudolinear Kalman filter that utilizes Euler's equations in the propagation of the estimated rate. Both algorithms are tested using flight data collected over a few months after the failure of two of the reaction wheels. The question of closed-loop stability is addressed. The ability of the controller to meet the science slew requirements, without the gyros, is analyzed.

An algorithm has been developed that converts penetration angles reported by Static Earth Sensors (SESs) into Earth observation vectors. This algorithm allows compensation for variation in the horizon height including that caused by Earth oblateness. It also allows pitch and roll to be computed using any number (greater than 1) of simultaneous sensor penetration angles simplifying processing during periods of Sun and Moon interference. The algorithm computes body frame unit vectors through each SES cluster. It also computes GCI vectors from the spacecraft to the position on the Earth's limb where each cluster detects the Earth's limb. These body frame vectors are used as sensor observation vectors and the GCI vectors are used as reference vectors in an attitude solution. The attitude, with the unobservable yaw discarded, is iteratively refined to provide the Earth observation vector solution.

The spin-stabilized Imager for Magnetopause to Aurora Global Exploration (IMAGE) is the National Aeronautics and Space Administration's (NASA's) first Medium-class Explorer Mission (MIDEX). IMAGE was launched into a highly elliptical polar orbit on March 25, 2000 from Vandenberg Air Force Base, California, aboard a Boeing Delta II 7326 launch vehicle. This paper presents some of the observations of the flight dynamics analyses during the launch and in-orbit checkout period through May 18, 2000. Three new algorithms - one algebraic and two differential correction - for computing the parameters of the coning motion of a spacecraft are described and evaluated using in-flight data from the autonomous star tracker (AST) on IMAGE. Other attitude aspects highlighted include support for active damping consequent upon the failure of the passive nutation damper, performance evaluation of the AST, evaluation of the Sun sensor and magnetometer using AST data, and magnetometer calibration.

This paper presents algorithms for estimating the angular-rate vector of satellites using quaternion measurements. Two approaches are compared one that uses differentiated quaternion measurements to yield coarse rate measurements, which are then fed into two different estimators. In the other approach the raw quaternion measurements themselves are fed directly into the two estimators. The two estimators rely on the ability to decompose the non-linear part of the rotas rotational dynamics equation of a body into a product of an angular-rate dependent matrix and the angular-rate vector itself. This non unique decomposition, enables the treatment of the nonlinear spacecraft (SC) dynamics model as a linear one and, thus, the application of a PseudoLinear Kalman Filter (PSELIKA). It also enables the application of a special Kalman filter which is based on the use of the solution of the State Dependent Algebraic Riccati Equation (SDARE) in order to compute the gain matrix and thus eliminates the need to compute recursively the filter covariance matrix. The replacement of the rotational dynamics by a simple Markov model is also examined. In this paper special consideration is given to the problem of delayed quaternion measurements. Two solutions to this problem are suggested and tested. Real Rossi X-Ray Timing Explorer (RXTE) data is used to test these algorithms, and results are presented.

this paper reports the results to date of early mission support provided by the personnel of the Goddard Space Flight Center Flight Dynamics Division (FDD) for the Rossi X-Ray Timing Explorer (RXTE) spacecraft. For this mission, the FDD supports onboard attitude determination and ephemeris propagation by supplying ground-based orbit and attitude solutions and calibration results. The first phase of that support was to provide launch window analyses. As the launch window was determined, acquisition attitudes were calculated and calibration slews were planned. postlaunch, these slews provided the basis for ground determined calibration. Ground determined calibration results are used to improve the accuracy of onboard solutions. The FDD is applying new calibration tools designed to facilitate use of the simultaneous, high-accuracy star observations from the two RXTE star trackers for ground attitude determination and calibration. An evaluation of the performance of these tools is presented. The FDD provides updates to the onboard star catalog based on preflight analysis and analysis of flight data. The in-flight results of the mission support in each area are summarized and compared with pre-mission expectations.

This paper characterizes the low-frequency noise response of the Teledyne dry rotor inertial reference unit (DRIRU) gyroscopes on the Upper Atmosphere Research Satellite (UARS) and the Extreme Ultraviolet Explorer (EUVE). The accuracy of spacecraft attitude estimation algorithms that use gyro data for propagating the spacecraft attitude is sensitive to gyro noise. EUVE gyro data were processed to validate a single-axis gyro noise model, which is used onboard various spacecraft. The paper addresses the potential impact of temperature effects on the gyro noise model and the overall impact on attitude determination accuracy. The power spectral density (PSD) of the gyro noise is estimated from UARS in-flight data by Fast Fourier Transform (FFT). The role of actuator dynamics on the PSD function is also discussed.

The Arthur Holly Compton Gamma Ray Observatory (GRO) was launched by the shuttle Atlantis in April 1991. This paper presents the results of the attitude sensor calibration that was performed during the early mission. The GSFC Flight Dynamics Facility (FDF) performed an alignment calibration of the two fixed-head star trackers (FHST's) and two fine Sun sensors (FSS's) on board Compton GRO. The results show a 27-arcsecond shift between the bore sights of the FHST's with respect to prelaunch measurements. The alignments of the two FSS's shifted by 0.20 and 0.05 degree. During the same time period, the Compton GRO science teams performed an alignment calibration of the science instruments with respect to the attitude reported by the on board computer (OBC). In order to preserve these science alignments, FDF adjusted the overall alignments of the FHST's and FSS's, obtained by the FDF calibration, such that when up linked to the OBC, the shift in the OBC-determined attitude is minimized. FDF also calibrated the inertial reference unit (IRU), which consists of three dual-axis gyroscopes. The observed gyro bias matched the bias that was solved for by the OBC. This bias drifted during the first 6 days after release. The results of the FDF calibration of scale factor and alignment shifts showed changes that were of the same order as their uncertainties.

The science mission of the Extreme Ultraviolet Explorer (EUVE) requires attitude solutions with uncertainties of 27, 16.7, 16.7 arcseconds (3 sigma) around the roll, pitch, and yaw axes, respectively. The primary input to the attitude determination process is provided by two NASA standard fixed-head star trackers (FHSTs) and a Teledyne dry rotor inertial reference unit (DRIRU) 2. The attitude determination requirements approach the limits attainable with the FHSTs and DRIRU. The Flight Dynamics Facility (FDF) at Goddard Space Flight Center (GSFC) designed and executed calibration procedures that far exceeded the extent and the data volume of any other FDF-supported mission. The techniques and results of this attempt to obtain attitude accuracies at the limit of sensor capability and the results of analysis of the factors that limit the attitude accuracy are the primary subjects of this paper. The success of the calibration effort is judged by the resulting measurement residuals and comparisons between ground- and onboard-determined attitudes. The FHST star position residuals have been reduced to less tha 4 arcsec per axis -- a value that appears to be limited by the sensor capabilities. The FDF ground system uses a batch least-squares estimator to determine attitude. The EUVE onboard computer (OBC) uses an extended Kalman filter. Currently, there are systematic differences between the two attitude solutions that occasionally exceed the mission requirements for 3 sigma attitude uncertainty. Attempts to understand and reduce these differences are continuing.

The accuracy of both onboard and ground attitude determination can be significantly enhanced by calibrating spacecraft attitude instruments (sensors) after launch. Although attitude sensors are accurately calibrated before launch, the stresses of launch and the space environment inevitably cause changes in sensor parameters. During the mission, these parameters may continue to drift requiring repeated on-orbit calibrations. The goal of attitude sensor calibration is to reduce the systematic errors in the measurement models. There are two stages at which systematic errors may enter. The first occurs in the conversion of sensor output into an observation vector in the sensor frame. The second occurs in the transformation of the vector from the sensor frame to the spacecraft attitude reference frame. This paper presents postlaunch alignment and transfer function calibration of the attitude sensors for the Compton Gamma Ray Observatory (GRO), the Upper Atmosphere Research Satellite (UARS), and the Extreme Ultraviolet Explorer (EUVE).

The results of on-orbit calibration for several satellites by the flight Dynamics Facility (FDF) at GSFC are reviewed. The examples discussed include attitude calibrations for sensors, including fixed-head star trackers, fine sun sensors, three-axis magnetometers, and inertial reference units taken from recent experience with the Compton Gamma Ray observatory, the Upper Atmosphere Research Satellite, and the Extreme Ultraviolet Explorer calibration. The methods used and the results of calibration are discussed, as are the improvements attained from in-flight calibration.

This analysis involves an in-depth look into the onboard computer (OBC) attitude determination algorithm. A review of TRW error analysis and necessary ground simulations to understand the onboard attitude determination process are performed. In addition, a plan is generated for the in-flight calibration and validation of OBC computed attitudes. Pre-mission expected accuracies are summarized and sensitivity of onboard algorithms to sensor anomalies and filter tuning parameters are addressed.

The requirements and mathematical specifications for the Gamma Ray Observatory (GRO) Dynamics Simulator are presented. The complete simulator system, which consists of the profie subsystem, simulation control and input/output subsystem, truth model subsystem, onboard computer model subsystem, and postprocessor, is described. The simulator will be used to evaluate and test the attitude determination and control models to be used on board GRO under conditions that simulate the expected in-flight environment.

This paper presents a covariance analysis of the performance and sensitivity of the attitude determination Extended Kalman Filter (EKF) used by the onboard computer (OBC) of the Extreme Ultra Violet Explorer (EUVE) spacecraft. The linearized dynamics and measurement equations of the error states are used in formulating the 'truth model' describing the order of the systems involved. The 'design model' used by the OBC EKF is then obtained by reducing the order of the truth model. The covariance matrix of the EKF which uses the reduced order model is not the correct covariance of the EKF estimation error. A 'true covariance analysis' has to be carried out in order to evaluate the correct accuracy of the OBC generated estimates. The results of such analysis are presented which indicate both the performance and the sensitivity of the OBC EKF.

This paper presents a covariance analysis of the performance and sensitivity of the attitude determination Extended Kalman Filter (EKF) used by the On Board Computer (OBC) of the Extreme Ultra Violet Explorer (EUVE) spacecraft. The linearized dynamics and measurement equations of the error states are used in formulating the 'truth model' describing the real behavior of the systems involved. The 'design model' used by the OBC EKF is then obtained by reducing the order of the truth model. The covariance matrix of the EKF which uses the reduced order model is not the correct covariance of the EKF estimation error. A 'true covariance analysis' has to be carried out in order to evaluate the correct accuracy of the OBC generated estimates. The results of such analysis are presented which indicate both the performance and the sensitivity of the OBC EKF.