Abstract: A spacecraft embedded in a reference frame rotating relative to inertial space. The spacecraft generally includes actuators for maneuvering the spacecraft with respect to the reference frame, an attitude measurement device that measures the pitch and roll attitude of the spacecraft with respect to the reference frame, a control device adapted to keep the roll and pitch angles of the spacecraft close to the commanded roll and pitch angles, and a harmonic torque estimator adapted to read the commanded angular velocity of the spacecraft relative to an inertial frame, read momentum wheel speeds, read known predicted external torques and combine angular velocity, measured wheel speed and known external torque to produce an estimated observable periodic torque.

Abstract: A spacecraft is provided that includes a communications payload having an antenna support structure and an antenna boom coupled to the antenna support structure. An attitude sensor may be mounted to the antenna boom. The attitude sensor (such as an Earth sensor, a star tracker and/or a sun sensor) may be mounted at a root of the antenna boom.

Abstract: A method for correcting differences in measurements between sensors. Misalignment of a roll error of a sun sensor alignment is estimated with respect to a reference roll value measured by an earth sensor. Sun sensor assembly elevation angle residual is calculated utilizing the estimated sun sensor assembly roll misalignment. A yaw attitude of the spacecraft is updated based upon the calculated sun sensor assembly elevation residual.

Abstract: By using the GPS satellite transmitted signals that are tracked on a satellite and those that should be tracked but are not due to the attitude of the satellite and rotating the satellite, two estimations of satellite attitude are made that are used to change the satellite attitude so that all the GPS signals are tracked.

Abstract: A spacecraft embedded in a reference frame rotating relative to inertial space. The spacecraft generally includes actuators for maneuvering the spacecraft with respect to the reference frame, an attitude measurement device that measures the pitch and roll attitude of the spacecraft with respect to the reference frame, a control device adapted to keep the roll and pitch angles of the spacecraft close to the commanded roll and pitch angles, and a harmonic torque estimator adapted to read the commanded angular velocity of the spacecraft relative to an inertial frame, read momentum wheel speeds, read known predicted external torques and combines angular velocity, measured wheel speed and known external torque to produce an estimated observable periodic torque.

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: The present invention provides methods of and apparatus for determining the inertial attitude of an aerospace vehicle. In one embodiment, the invention provides a rotational astronomical object-sighting concept to determine the inertial attitude of an axis of the aerospace vehicle without the star identification or dragback. In another embodiment, the invention provides an attitude measurement apparatus comprising a high sensitivity optical sensor and a low power inertial sensor.

Abstract: A simple, robust algorithm for power acquisition for power, thermal, and momentum safety for high heater-power spacecraft uses current sensors rather than sun sensors and includes a wing sun search phase, an xz slew phase, and a safe hold phase. Solar wing current is continuously monitored against a high current threshold and a low current threshold. Wing sun search phase transitions either to xz slew phase or safe hold phase. When current is too low, the xz slew phase is entered and slews the spacecraft about an axis in the xz plane until current is high enough. When current is high enough, the algorithm transitions or remains in safe hold phase and the spacecraft is spun about its wing axis. A low current persistence timer is longer than a high current persistence timer in order to bias the algorithm to prefer safe hold phase over xz slew phase.

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 spacecraft power/sun acquisition method includes a pitch search phase in which the spacecraft is rotated about a pitch axis, stopped, quaternions reset to remember position, and yaw search phase is entered when a first TOA occurs from a first slit sun sensor, and a keyhole slew is performed when the first TOA does not occur. The method also includes yaw search phase in which a keyhole slew is performed by rotating the spacecraft about a yaw axis, stopping the spacecraft, and entering pitch search phase when the first TOA does not occur, and in which the spacecraft is rotated about a yaw axis, stopped, quaternions reset, and sun hold phase is entered when a second TOA occurs from a second slit sun sensor. Lastly, the method includes a sun hold phase in which the spacecraft is oriented to the sun and placed in spin at power safe attitude.

Abstract: An apparatus for determining the attitude of an artificial satellite by narrowing attitude candidates to a correct one in a short time. The apparatus has star sensors, a star catalog data base, star identification sections in each of which star images output from the star sensors are collated with a star catalog with respect to each star sensor to output a group of corresponding candidates, attitude computation sections for computing a value of an attitude candidate of the artificial satellite with respect to each corresponding candidate, an attitude updating section for updating the value of the attitude candidate on the basis of the star images output from the star sensor and the star catalog, and an attitude propagation section for computing the value of the attitude candidate at the present sampling time from the value of the attitude candidate at the preceding sampling time and the artificial satellite's angular velocity.

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 satellite attitude control system includes a programmed processor system which includes a gyroscopic actuator first control stage for changing the attitude of the satellite and a reaction wheel second control stage for assuring that pointing of the satellite is accurate and stable. The method is intended to be used for a satellite including the two control stages indicated above, which it uses selectively for the operations indicated above.

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: An apparatus, method and computer program product useful for supplying the pilot of an aircraft with aircraft attitude information in the event of full or partial failure of the gyroscopic system normally used to supply such data. The pitch angle, roll angle and heading angle estimates provided by the apparatus, method and computer program product can be displayed to the pilot thereby alleviating the requirement that the pilot mentally integrate such data from the remaining aircraft instruments.

Abstract: An attitude control system uses a first order partial derivative of an error signal, representing a difference between a commanded pitch attitude and a sensed, actual pitch attitude. The first order partial derivative forms a cross product with the Earth's magnetic field to provide a X axis torque rod dipole command, which powers the torque rod in a way that causes a satellite to move toward its commanded attitude.

Abstract: A system and method for correcting pointing errors induced by an inclined orbit (14) for a satellite (20) that is adapted to be in an orbit having zero inclination (12). The present invention generates a yaw correcting cosine rate command (34) that is integrated (36) by a steering law having input parameters dependent upon the angle of inclination, &THgr;, to derive a sinusoidal yaw offset. The present invention also generates a roll correcting cosine rate command (54) that is integrated (56) according to a steering law having input parameters dependent upon the angle of inclination, &THgr;, to derive a sinusoidal roll offset.

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: This invention provides an offline attitude determining apparatus capable of detecting an artificial satellite attitude at a high precision in wide band. Averaging and addition processing are carried out on a low-frequency attitude interpolation signal generated by interpolating a low-frequency attitude detection signal obtained by inputting an attitude angle determination signal generated by a sequential Kalman filter computation device and a high-frequency attitude detection signal obtained by inputting a high-frequency attitude angle sensor detection signal extracted by a high-frequency attitude angle sensor data extracting device into a band pass filter computation device, so as to generate a high-precision attitude detection signal. By combining the low-frequency attitude detection signal with the high-frequency attitude detection signal with an interpolation processing computation device and an attitude data generator, the attitude can be determined at a high precision in wide band.

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: A process for determining an attitude measurement of a spacecraft is provided. The process employs a Horizon-Crossing Indicator (HCI) sensor 108. The process uses three time tagged crossings of the earth circle 104 to provide sufficient HCI data to determine the earth center 106. Since these observations may be infrequent, the relative attitude of the spacecraft 100 changes for each observation time. Therefore, once the integrated angular rate data, and thus the attitude relative to some reference, for the first, second and third horizon crossings is obtained, the sensor bore-sight data for each of these observations is transformed back to a known reference time. In one example, the integrated angular rate data from the third and second horizon crossings is transformed back to first horizon crossing time T1.

Abstract: A control system for providing attitude control in spacecraft. The control system comprising a primary attitude reference system, a secondary attitude reference system, and a hyper-complex number differencing system. The hyper-complex number differencing system is connectable to the primary attitude reference system and the secondary attitude reference system.

Abstract: An optical sensor includes dual fields of view including a panoramic field of view spanning 360° in azimuth angle in a direction perpendicular to an axis of the sensor, and a limb-looking field of view non-perpendicular to the axis for viewing the limb of the earth. Both fields of view are imaged onto annular regions of one of more focal plane arrays comprising pixels arranged in a rectangular array of rows and columns. The sensor is used in a method for capturing rate and direction of rotation of a satellite about its axes and for detecting orientation of the satellite about two of its axes. Rate and direction are determined by finding the center of the earth relative to axes of the focal plane array based on the image of the earth limb from the panoramic field of view, and comparing the earth center location at a series of sequential times. The rate and position information are used for stabilizing an initially tumbling satellite after tip-off.

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 optical sensor includes dual fields of view including a panoramic field of view spanning 360° in azimuth angle in a direction perpendicular to an axis of the sensor, and a limb-looking field of view non-perpendicular to the axis for viewing the limb of the earth. Both fields of view are imaged onto annular regions of one of more focal plane arrays comprising pixels arranged in a rectangular array of rows and columns. The sensor is used in a method for capturing rate and direction of rotation of a satellite about its axes and for detecting orientation of the satellite about two of its axes. Rate and direction are determined by finding the center of the earth relative to axes of the focal plane array based on the image of the earth limb from the panoramic field of view, and comparing the earth center location at a series of sequential times. The rate and position information are used for stabilizing an initially tumbling satellite after tip-off.

Abstract: A method, apparatus, and article of manufacture for directing a payload coupled to a spacecraft substantially at a target location is disclosed. The method comprises the steps of determining spacecraft orbital parameters and spacecraft payload attitude parameters; and directing the payload at the target location by applying a spacecraft bus steering law determined at least in part from the orbit and attitude parameters. The apparatus comprises a navigation system for determining spacecraft orbit parameters and spacecraft attitude parameters, and for applying a spacecraft bus steering law to direct the payload substantially at the target location. The spacecraft bus steering law is determined at least in part from the determined spacecraft orbital parameters, the determined spacecraft attitude, and determined payload attitude parameters. In one embodiment, the apparatus includes a payload attitude control system for directing the payload relative to the spacecraft bus.

Abstract: Positioning system for a high-precision measuring instrument (1) in a support structure (2) of a satellite, in which the measuring instrument (1) can be moved freely with predetermined degrees of freedom within a space (3) provided in the satellite support structure (2). The measuring instrument (1) produces reference signals required by it for its inertial alignment, and is aligned with respect to a predetermined measurement target by a first positioning device (7). The satellite support structure (2) follows the motion of the measuring instrument (1) in order to overcome external disruptive forces and moments.

Abstract: The attitude of a tiltable body is tracked and controlled on the basis of signals outputted from gyros and tilt sensors. The signals outputted from the gyros are transformed and integrated to produce estimated position information in the form of a modified quaternion in which the yaw component is constrained to a zero value. Modified quaternion information in the same form is also generated from signals outputted from the tilt sensor and is used to detect and correct an error component in the estimated position information. Gyro drift is also corrected based on the output of the tilt sensor.

Abstract: This invention provides an offline attitude determining apparatus capable of detecting an artificial satellite attitude at a high precision in wide band. Averaging and addition processing are carried out on a low-frequency attitude interpolation signal generated by interpolating a low-frequency attitude detection signal obtained by inputting an attitude angle determination signal generated by a sequential Kalman filter computation device and a high-frequency attitude detection signal obtained by inputting a high-frequency attitude angle sensor detection signal extracted by a high-frequency attitude angle sensor data extracting device into a band pass filter computation device, so as to generate a high-precision attitude detection signal. By combining the low-frequency attitude detection signal with the high-frequency attitude detection signal with an interpolation processing computation device and an attitude data generator, the attitude can be determined at a high precision in wide band.

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.

Abstract: A microaltimeter measures altitude or range highly accurately from an orbiting vehicle. The microaltimeter has a low power solid state laser that is pulsed at a rate above 1 kilohertz. The pulses are delivered to a small telescope which sends them to a planetary surface and receives return reflections. A high efficiency photon detector measures received photons and supplies received photon signals to a process or which makes a time-based bin-wise comparison to find the time of flight and hence the range.

Abstract: An apparatus for star catalog equalization to enhance attitude determination 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 star location data stored in the star catalog is predetermined by calculating a plurality of desired star locations, associating one of a plurality of stars with each of the plurality of desired star locations based upon a neighborhood association angle to generate an associated plurality of star locations: If an artificial star gap occurs during association, then the neighborhood association angle for reassociation is increased.

Abstract: A method for use on a satellite that automatically inhibits scanning of an Earth sensor to handle sensor intrusions by the sun, moon, or other celestial bodies. In implementing the method, a predicted state vector for the satellite, derived from an orbit propagator, is generated. An attitude profile for the satellite is generated. Then, the satellite state (predicted state vector) and attitude profile are processed to determine Earth, sun, and moon vectors in a satellite body frame of reference at any instant. The Earth, sun, and moon vectors are compared to the Earth sensor field of view and sensor field of view limit boxes to determine if scan inhibiting or Earth sensor switching should occur. The affected Earth sensor is inhibited or switched if an intrusion of the sun and/or moon into the field of view of the Earth sensor is predicted.

Abstract: An apparatus, method and computer program product useful for supplying the pilot of an aircraft with aircraft attitude information in the event of full or partial failure of the gyroscopic system normally used to supply such data. The pitch angle, roll angle and heading angle estimates provided by the apparatus, method and computer program product can be displayed to the pilot thereby alleviating the requirement that the pilot mentally integrate such data from the remaining aircraft instruments.

Abstract: A spacecraft having a body, one or more appendages coupled thereto, and a controller that implements methods that rotate the one or more flexible appendages to point it (them) towards the Sun to reduce spacecraft attitude pointing disturbances and improves spacecraft attitude pointing. The steps of the one or more appendages are timed to deadbeat the disturbance imparted to the spacecraft body. Timing of the appendage steps may be such that the periodic disturbances are phased to substantially cancel each other, or phased to decrease the magnitude of the net disturbance. The present invention also cancels solar array slew disturbances. The present invention cancels predictable disturbance torques before they produce a pointing error, improving the spacecraft pointing performance.

Abstract: The present invention is embodied in an apparatus for sensing solar position. The apparatus includes a precision wide-field-of-view (WFOV) passive optical sensor head and a set of interface electronics disposed within a spacecraft computer, which is located remotely from the precision WFOV optical sensor head. The interface electronics convert current signals supplied by the precision WFOV optical sensor head into digital signals usable by a spacecraft control processor. The spacecraft control processor uses the digital signals from the interface electronics to precisely determine the attitude of the spacecraft.

Abstract: A system and method of estimating the attitude of a spacecraft is disclosed. A three-axis inertial-based estimate of spacecraft attitude is produced and is compared to a stellar-based estimate of spacecraft attitude. A Kalman filter having states associated with gyro scale factor and/or misalignment errors compares the stellar-based attitude estimate to the inertial-based estimate of attitude and apportions the total error into three time varying matrices. A first time varying matrix is associated with gyro scale factor and misalignment errors, a second time varying matrix is associated with gyro bias errors, and a third time varying matrix is associated with attitude errors. The time varying matrices are applied as corrective feedbacks to the inertial-based estimate of spacecraft attitude and are adaptively adjusted to minimize the error therein.

Abstract: A momentum control system for driving the momentum of a spinning spacecraft to zero while it spins around a sunline. The system includes a sun sensor for determining the position of the sun, a suite of tachometers for determining the momentum stored in the spacecraft, and a stored movement sensor. Each of the readings from these sensors is fed to a controller. The controller also monitors the magnitude of the spacecraft overturning momentum and the windmill momentum. If the overturning momentum is outside acceptable limits then the rotation rate of the spacecraft is modulated to drive the overturning momentum towards zero. If the windmill momentum exceeds acceptable limits, the angle of the solar panel is adjusted to drive the windmill momentum towards zero.

Abstract: Techniques for monitoring and controlling yawing of a GPS satellite in an orbit that has an eclipsing portion out of the sunlight based on the orbital conditions of the GPS satellite. In one embodiment, a constant yaw bias is generated in the attitude control system of the GPS satellite to control the yawing of the GPS satellite when it is in the shadow of the earth.

Abstract: A star tracker coupled to the spacecraft having a star catalog associated therewith. A sun sensor is coupled to the spacecraft. A control processor is coupled to the star tracker and the sun sensor. The processor obtains star data using a star tracker and an on-board star catalog. The processor generates a coarse attitude of the spacecraft as a function of the star data, and establishes a track on at least one star in the on-board star catalog. The processor calculates a sun tracking rate, and obtains a normal phase attitude as a function of the star data and the coarse attitude. The information is used to slew the spacecraft to a desired attitude.

Abstract: A method and system for minimizing attitude errors resulting from dynamic spacecraft maneuvers or time-varying mismatched attitude sensor measurements. The method includes time-tagging a primary attitude value from a primary attitude sensor and storing the value and associated time stamp in a buffer. A secondary attitude value, generated by a secondary attitude sensor, having a different associated time-stamp is then time-matched to the primary attitude valve. The corrected spacecraft attitude is then calculated as a function of the time-matched primary and secondary attitude values. In one aspect of the invention, the primary attitude sensor generates attitude data at a faster rate than the secondary attitude sensor. This primary attitude data is stored in a buffer such that, when the secondary attitude data becomes available, several nearest-in-time primary attitude data points are retrieved from the buffer. These data points are then used to interpolate a higher time-matched resolution data point.

Abstract: A smart docking surface consisting of closely spaced cantilevered sensor/actuator structures capable of precisely repositioning an object having a ferro-magnetic surface in contact with the smart docking surface. It is designed for use in a micro gravity environment for the final docking sequence of two small (<100 kg) satellites. Its purpose is to reduce the complexity of the docking process where a precise mating is required.

Abstract: A system and method of estimating the attitude of a spacecraft compares a three-axis inertial-based estimate of spacecraft attitude to a stellar-based estimate of the spacecraft attitude. A Kalman filter having some states associated with low spatial frequency errors compares the stellar-based attitude estimate to the inertial-based estimate of attitude and apportions total attitude error into two time varying matrices. A first time varying matrix is associated with star tracker low spatial frequency errors, a second time varying matrix is associated with gyro bias errors and attitude errors. The time varying matrices are used to apply corrective feedbacks to the stellar-based attitude estimate and the inertial-based estimate of spacecraft attitude, and are adaptively adjusted to minimize total estimated attitude error.

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 preparing, for each star sensor, an associated star catalog having an entry describing star parameters for each star potentially observable by the star sensor. The determination of which stars are potentially observable by the star sensor is performed using an estimate of the satellite attitude. A correlation is then determined between the stars observed by each star sensor with the associated star catalog entries. A reference star sensor is selected as the star sensor tracking a primary star having the highest correlation between the stars observed by the star sensor and the entries in the star catalog associated with the star sensor. The spacecraft attitude is determined from the primary star and remaining star observations by the reference star sensor. Attitude estimates from the non-reference trackers are used to determine star sensor alignment.

Abstract: A pseudo gyro emulates mechanical gyros by software processes by processing space system appendage measurement data and reaction wheel tachometer data within reference and control systems of a satellite using principals of conservation of momentum to compute the vehicular bus angular velocity rate data by accounting for external torque, the momentum transfer between the satellite, the bus, and the appendages for providing accurate relative vehicular position and angular velocity rate data as an integral part of attitude reference and control systems now having higher reliability, longer life times, lower power consumption, and more accurate vehicular angular velocity rate data generated within high bandwidth operations. A specific improvement is the use of a bias torque estimator for accounting for uncertainties in external torque estimations for accurate total system momentum computations.

Abstract: High precision spacecraft attitude determination is produced by specially positioning the spacecraft's star trackers and then filtering out measurement errors produced from star tracker electronics. In addition to the conventional azimuth and elevation controls used for star tracker pointing, the star tracker detector array is rotationally positioned about its boresight so that its pixels are traversed by the imaged star path at an angle within 20° of diagonal. This forces both vertical and horizontal spatial error components in the detector plane to a high frequency range at which they can easily be filtered out in common.

Abstract: An attitude determination system for an artificial satellite capable of performing star identification without need for the aid of ground station includes an image processing module (17) for processing star images observed at predetermined time points by a star sensor (16) mounted on the artificial satellite (1) for arithmetically determining direction vectors of the observed stars, a rotation estimating module (18) for estimating a rotational motion of the artificial satellite (1) between an attitude of the artificial satellite at a predetermined time point and an attitude of the artificial satellite at another time point, an elongation estimating module (19) for estimating elongations between the direction vectors of plural stars the images of which are picked up at a same time point by the star sensor and estimating the elongations between the direction vectors of plural stars the images of which are picked up at different time points by the star sensor on the basis of the estimated rotational motion, a st