Abstract: Methods and structures are provided that enhance attitude control during gyroscope substitutions by insuring that a spacecraft's attitude control system does not drive its absolute-attitude sensors out of their capture ranges. In a method embodiment, an operational process-noise covariance Q of a Kalman filter is temporarily replaced with a substantially greater interim process-noise covariance Q. This replacement increases the weight given to the most recent attitude measurements and hastens the reduction of attitude errors and gyroscope bias errors. The error effect of the substituted gyroscopes is reduced and the absolute-attitude sensors are not driven out of their capture range. In another method embodiment, this replacement is preceded by the temporary replacement of an operational measurement-noise variance R with a substantially larger interim measurement-noise variance R to reduce transients during the gyroscope substitutions.

Abstract: Methods and structures are provided for reducing pointing errors &zgr; of satellite antennas and for generating broad field-of-view satellite attitude acquisition patterns. In one method embodiment, satellite transmit beams have estimated pointing attitudes &bgr; and are transmitted to overlap on a ground-based receiving terminal which has a known terminal location &lgr; and which measures received signal strengths &agr;. Pointing errors &zgr; of the transmit beams are then determined from the estimated pointing attitudes &bgr;, the terminal location &lgr; and the signal strengths &agr; and the pointing errors &zgr; are subsequently reduced by revising the pointing attitudes &bgr;. Other method embodiments utilize known signal-strength functions and antenna signals with known signal parameters such as frequencies and/or modulations.

Abstract: Methods and structures are provided that enhance attitude control during gyroscope substitutions by insuring that a spacecraft's attitude control system does not drive its absolute-attitude sensors out of their capture ranges. In a method embodiment, an operational process-noise covariance Q of a Kalman filter is temporarily replaced with a substantially greater interim process-noise covariance Q. This replacement increases the weight given to the most recent attitude measurements and hastens the reduction of attitude errors and gyroscope bias errors. The error effect of the substituted gyroscopes is reduced and the absolute-attitude sensors are not driven out of their capture range. In another method embodiment, this replacement is preceded by the temporary replacement of an operational measurement-noise variance R with a substantially larger interim measurement-noise variance R to reduce transients during the gyroscope substitutions.

Abstract: A method and apparatus for calibrating rate sensor measurements to compensate for rate sensor scale factor non-linearities. An outer loop compensator compares current rate sensor scale factor estimates with the current scale factor non-linearity compensation, and deviations from the current non-linearity compensation are corrected in an updated compensation.

Abstract: Attitude determination and control systems are provided that combine attitude measurements from all spacecraft payloads to determine a master attitude estimate for a master payload and relative slave attitude estimates for the remaining slave payloads. These estimates are then used to control the attitudes of spacecraft elements that correct the absolute and relative attitude errors. These systems significantly enhance attitude accuracy when compared to systems that realize independent payload estimates, determine payload attitudes. These systems also provide significant processing advantages (e.g., simpler algorithms, reduced data throughput and slower processing rate).

Abstract: The present invention is directed to spacecraft that have, for any reason, lost the spacecraft's service attitude that permits it to carry out the service operations for which it was designed. The invention provides methods and structures for acquiring and determining a power-safe attitude (i.e., one in which wing current is sufficient to support the spacecraft's housekeeping operations) from which the spacecraft can be subsequently returned to a service attitude. The methods are particularly useful because they a) require only a single star tracker for sensing attitude, comprise simple maneuvers, and typically acquire a power-safe attitude that does not significantly differ from the spacecraft's service attitude to thereby reduce the spacecraft's return-to-service time.

Abstract: A method and apparatus for post processing a star tracker measurement to remove a systematic error characterizable at least in part by a pixel phase is disclosed. The method comprises the steps of computing the pixel phase along a first axis from a measured star position and a star tracker characteristic, computing a first axis error correction according to the computed pixel phase, and computing a compensated first axis star tracker measurement according to the measured star position and the first axis error correction.

Abstract: A method and apparatus for calibrating rate sensor measurements to compensate for rate sensor scale factor non-linearities. An outer loop compensator compares current rate sensor scale factor estimates with the current scale factor non-linearity compensation, and deviations from the current non-linearity compensation are corrected in an updated compensation.

Abstract: Attitude determination and control systems are provided that combine attitude measurements from all spacecraft payloads to determine a master attitude estimate for a master payload and relative slave attitude estimates for the remaining slave payloads. These estimates are then used to control the attitudes of spacecraft elements that correct the absolute and relative attitude errors. These systems significantly enhance attitude accuracy when compared to systems that realize independent payload estimates. determine payload attitudes. These systems also provide significant processing advantages (e.g., simpler algorithms, reduced data throughput and slower processing rate).

Abstract: A method and apparatus for post processing a star tracker measurement to remove a systematic error characterizable at least in part by a pixel phase is disclosed. The method comprises the steps of computing the pixel phase along a first axis from a measured star position and a star tracker characteristic, computing a first axis error correction according to the computed pixel phase, and computing a compensated first axis star tracker measurement according to the measured star position and the first axis error correction.

Abstract: The present invention is directed to spacecraft that have, for any reason, lost the spacecraft's service attitude that permits it to carry out the service operations for which it was designed. The invention provides methods and structures for acquiring and determining a power-safe attitude (i.e., one in which wing current is sufficient to support the spacecraft's housekeeping operations) from which the spacecraft can be subsequently returned to a service attitude. The methods are particularly useful because they a) require only a single star tracker for sensing attitude, comprise simple maneuvers, and typically acquire a power-safe attitude that does not significantly differ from the spacecraft's service attitude to thereby reduce the spacecraft's return-to-service time.

Abstract: A method, apparatus, article of manufacture for determining the attitude of a spacecraft having at least one star sensor. The method comprises the steps of selecting a first reference star sensor from among the star sensors; designating two stars observed by the first reference star sensor as a primary star pair; identifying a candidate star pair corresponding to the primary star pair, wherein the candidate star pair is selected from a star catalog having a plurality of entries; estimating an inertial orientation of the first star sensor at least in part from the identified candidate star pair; and determining the spacecraft attitude from the estimated inertial orientation of a reference star sensor selected from a group comprising the first star sensor.

Abstract: A method, apparatus, article of manufacture for determining the attitude of a spacecraft having at least one star sensor. The method comprises the steps of selecting a first reference star sensor from among the star sensors; designating two stars observed by the first reference star sensor as a primary star pair; identifying a candidate star pair corresponding to the primary star pair, wherein the candidate star pair is selected from a star catalog having a plurality of entries; estimating an inertial orientation of the first star sensor at least in part from the identified candidate star pair; and determining the spacecraft attitude from the estimated inertial orientation of a reference star sensor selected from a group comprising the first star sensor.

Abstract: A pointing system and method for directing a first satellite system instrument to a first target is disclosed. The method comprises the steps of computing the orientation of a first satellite system instrument; computing a position of the satellite using measured orbital data; computing a target line of sight (LOS) vector direction from the satellite to the target using the computed orientation and the computed satellite position, and directing the first instrument to the first target according to the computed target LOS vector. The apparatus comprises at least one attitude sensor for the first instrument and a communicatively coupled satellite navigation system. The satellite navigation system determines the satellite position using measured satellite orbital data. The satellite navigation system computes the orientation of the first instrument using the attitude measurements, and computes the satellite position using measured satellite orbital data.

Abstract: An apparatus for determining star location includes a star tracker, a star catalog and a controller. The star tracker is used to sense the positions of stars and generate signals corresponding to the positions of the stars as seen in its field of view. The star catalog contains star location data that is stored using a primary and multiple secondary arrays sorted by both declination (DEC) and right ascension (RA), respectively. The controller checks the star catalog and determines which stars to track. The controller does this determination by using an algorithm to sort the primary and secondary arrays to determine which stars are located in the star tracker field of view. The controller then commands the star tracker to track these stars and uses them to determine the spacecraft attitude.

Abstract: A method, apparatus, article of manufacture for determining the attitude of a spacecraft having at least one star sensor. The method comprises the steps of selecting a first reference star sensor from among the star sensors; designating two stars observed by the first reference star sensor as a primary star pair; identifying a candidate star pair corresponding to the primary star pair, wherein the candidate star pair is selected from a star catalog having a plurality of entries; estimating an inertial orientation of the first star sensor at least in part from the identified candidate star pair; and determining the spacecraft attitude from the estimated inertial orientation of a reference star sensor selected from a group comprising the first star sensor.

Abstract: Structures and methods are provided for deriving corrected star coordinates Ccrctd from measured star coordinates Cms that include star tracker charge transfer efficiency (CTE) errors. The structures and methods are based on a recognition that measured star coordinates Cms of star image centroids include CTE errors which are functions of the CCD path lengths over which the associated electrical charges traveled. In particular, the errors are substantially a product of a respective path length and a star-coordinate error factor &xgr; which, in turn, is a function of the star image magnitudes msi. Information contained in different measured star coordinates Cms is organized to facilitate the derivation of an estimate &xgr;* of the star-coordinate error factor &xgr; with conventional estimation processes. The measured star coordinates Cms are then corrected with the error factor estimate &xgr;* to realize the corrected star coordinates Ccrctd and, thereby, improve the accuracy of spacecraft attitude control.

Abstract: Structures and methods are provided for deriving corrected star coordinates Ccrctd from measured star coordinates Cms that include star tracker charge transfer efficiency (CTE) errors. The structures and methods are based on a recognition that measured star coordinates Cms of star image centroids include CTE errors which are functions of the CCD path lengths over which the associated electrical charges traveled. In particular, the errors are substantially a product of a respective path length and a star-coordinate error factor &xgr; which, in turn, is a function of the star image magnitudes msi. Information contained in different measured star coordinates Cms is organized to facilitate the derivation of an estimate &xgr;* of the star-coordinate error factor &xgr; with conventional estimation processes. The measured star coordinates Cms are then corrected with the error factor estimate &xgr;* to realize the corrected star coordinates Ccrctd and, thereby, improve the accuracy of spacecraft attitude control.

Abstract: An apparatus for determining stellar inertial attitude based upon a plurality of stars includes a star tracker and a controller. The star tracker is used to sense the positions of stars and generate signals corresponding to the positions of the stars as seen in its field of view. The controller checks a star catalog and actively determines which stars to track. The controller does this determination in terms of each star's relative geometry. The controller then commands the star tracker to track these stars and uses them to determine the spacecraft attitude.

Abstract: Stellar attitude-control systems and methods are provided with enhanced accuracy because they recognize that important star tracker errors exhibit a boresight symmetry and that these errors can be accurately defined by weighting a measurement-noise covariance matrix R(tn) with variances that are functions of off-boresight angles of detected stars. A method of the invention derives off-boresight angles &thgr; from star-tracker signals of detected stars. These off-boresight angles &thgr; are combined with variance coefficients &agr; to generate off-boresight variances rob(tn) that are functions of the off-boresight angles &thgr; and, in particular, correspond to star tracker color shift errors. A gain matrix K(tn) is then calculated with a weighted measurement-noise covariance matrix R(tn) that includes the off-boresight variances rob(tn). This gain matrix is used to generate an attitude estimate matrix X*(tn) for use in spacecraft attitude control.