Abstract: Sensor data fusion systems that provide noise reduction and fault protection. The sensor data fusion system fuses data acquired by respective accelerometers having different attributes. For example, one accelerometer has low noise and high bias, while another accelerometer has high noise and low bias when measuring specific force. The high-noise, low-bias accelerometer may be a gravimeter. Gravimeters and traditional accelerometers measure the same physical variable, i.e., specific force. By combining an expensive gravimeter and low-cost accelerometers, a synthetic sensor having both low noise and low bias may be achieved. Such synthetic sensors may be utilized in a gravity anomaly-referenced navigation system to achieve improved navigation performance.

Abstract: A celestial navigation system and method for determining a position of a vehicle. The system includes a star-tracker, a beam director, an inertial measurement unit, and a control module. The star-tracker has a field of view for capturing light. The beam director is configured to change a direction of the light captured in the field of view of the star-tracker. The inertial measurement unit has a plurality of sensors for measuring an acceleration and a rotation rate of the vehicle. The control module executes instructions to correct the attitude, the velocity and the position of the vehicle using the determined magnitude and position of the space objects. The control module also executes instructions to generate corrections to the IMU error parameters, the beam director and star-tracker alignment errors, and RSO ephemeris errors to achieve optimal performance.

Abstract: A celestial navigation system and method for determining a position of a vehicle. The system includes a star-tracker, a beam director, an inertial measurement unit, and a control module. The star-tracker has a field of view for capturing light. The beam director is configured to change a direction of the light captured in the field of view of the star-tracker. The inertial measurement unit has a plurality of sensors for measuring an acceleration and a rotation rate of the vehicle. The control module executes instructions to correct the attitude, the velocity and the position of the vehicle using the determined magnitude and position of the space objects. The control module also executes instructions to generate corrections to the IMU error parameters, the beam director and star-tracker alignment errors, and RSO ephemeris errors to achieve optimal performance.

Abstract: Systems and methods for providing improved navigation performance in which camera images are matched (using correlation) against reference images available from a geolocation-tagged database. An image-based navigation system partitions the camera image (corresponding to the area within the field-of-view of the camera) into a plurality of camera sub-images (corresponding to regions in that area), and then further partitions each camera sub-image into a multiplicity of tiles (corresponding to sub-regions within a region). The partitioning into camera sub-images seeks geometric diversity in the landscape. Each tile is checked for quality assurance, including feature richness, before correlation is attempted. The correlation results are further quality-checked/controlled before the results are used by the Kalman filter to generate corrections for use by the inertial navigation system.

Abstract: An attitude estimator that uses sun sensor outputs as the only attitude determination measurements to provide three-axis attitude information. This is accomplished by incorporating the Euler equation into the estimator. An unscented Kalman filter is employed to accommodate various nonlinear characteristics and uncertainties of the spacecraft dynamics and thus improve the robustness and accuracy of the attitude estimate.

Abstract: An attitude estimator that uses star tracker measurements and enhanced Kalman filtering, with or without attitude data, to provide three-axis rate estimates. The enhanced Kalman filtering comprises taking an average of forward and rearward propagations of the Kalman filter states and the error covariances. The star tracker-based rate estimates can be used to control the attitude of a satellite or to calibrate a sensor, such as a gyroscope.

Abstract: An attitude estimator that uses star tracker measurements and enhanced Kalman filtering, with or without attitude data, to provide three-axis rate estimates. The enhanced Kalman filtering comprises taking an average of forward and rearward propagations of the Kalman filter states and the error covariances. The star tracker-based rate estimates can be used to control the attitude of a satellite or to calibrate a sensor, such as a gyroscope.

Abstract: An attitude estimator that uses sun sensor outputs as the only attitude determination measurements to provide three-axis attitude information. This is accomplished by incorporating the Euler equation into the estimator. An unscented Kalman filter is employed to accommodate various nonlinear characteristics and uncertainties of the spacecraft dynamics and thus improve the robustness and accuracy of the attitude estimate.

Abstract: Methods, systems, and computer-readable media are described herein for using a modified Kalman filter to generate attitude error corrections. Attitude measurements are received from primary and secondary attitude sensors of a satellite or other spacecraft. Attitude error correction values for the attitude measurements from the primary attitude sensors are calculated based on the attitude measurements from the secondary attitude sensors using expanded equations derived for a subset of a plurality of block sub-matrices partitioned from the matrices of a Kalman filter, with the remaining of the plurality of block sub-matrices being pre-calculated and programmed into a flight computer of the spacecraft. The propagation of covariance is accomplished via a single step execution of the method irrespective of the secondary attitude sensor measurement period.

Abstract: Methods, systems, and computer-readable media are described herein for using a modified Kalman filter to generate attitude error corrections. Attitude measurements are received from primary and secondary attitude sensors of a satellite or other spacecraft. Attitude error correction values for the attitude measurements from the primary attitude sensors are calculated based on the attitude measurements from the secondary attitude sensors using expanded equations derived for a subset of a plurality of block sub-matrices partitioned from the matrices of a Kalman filter, with the remaining of the plurality of block sub-matrices being pre-calculated and programmed into a flight computer of the spacecraft. The propagation of covariance is accomplished via a single step execution of the method irrespective of the secondary attitude sensor measurement period.

Abstract: A method for controlling a vehicle may include sensing a position of each of a plurality of stars relative to the vehicle. The method may also include determining an attitude of the vehicle using the sensed positions of the plurality of stars, and the attitude may be determined either with or without using information from a gyro or sensor for measuring angular velocity. The method may additionally include implementing a set of strategies to optimize determination of the attitude of the vehicle when using only the sensed positions of the plurality of stars, without information from the sensor for measuring angular velocity. The method may further include controlling the vehicle based on the determined attitude of the vehicle.

Abstract: A system and method for gyroless transfer orbit sun acquisition using only wing current measurement feedback is disclosed. With this system and method, a spacecraft is able to maneuver itself to orient its solar panel to its maximum solar exposure spinning attitude. The disclosed system and method involve controlling a spacecraft maneuver using only the solar wing current feedback as the sole closed-loop feedback sensor for attitude control. A spin controller is used for controlling the spacecraft spin axis orientation and spin rate. The spin controller commands the spacecraft spin axis orientation to align with an inertial fixed-direction and to rotate at a specified spin rate by using a momentum vector. In addition, a method for estimating spacecraft body angular rate and spacecraft attitude is disclosed. This method uses a combination of solar array current and spacecraft momentum as the cost function with solar wing current feedback as the only closed-loop feedback sensor.

Abstract: A system and method for gyroless transfer orbit sun acquisition using only wing current measurement feedback is disclosed. With this system and method, a spacecraft is able to maneuver itself to orient its solar panel to its maximum solar exposure spinning attitude. The disclosed system and method involve controlling a spacecraft maneuver using only the solar wing current feedback as the sole closed-loop feedback sensor for attitude control. A spin controller is used for controlling the spacecraft spin axis orientation and spin rate. The spin controller commands the spacecraft spin axis orientation to align with an inertial fixed-direction and to rotate at a specified spin rate by using a momentum vector. In addition, a method for estimating spacecraft body angular rate and spacecraft attitude is disclosed. This method uses a combination of solar array current and spacecraft momentum as the cost function with solar wing current feedback as the only closed-loop feedback sensor.

Abstract: A method and apparatus for controlling a gimbaled platform (108). The method comprises the steps of computing an acquisition phase gimbal angle rate command ?cmd—Acq (530) from a measured LOS angle error ??LOS (520) for an initial control period T while computing an estimated LOS angle rate {circumflex over (?)}LOS (524), computing a tracking phase gimbal angle rate command ?cmd—Trk (532) using a controller (512) having an output initialized with the estimated LOS angle rate {circumflex over (?)}LOS (524), and commanding the gimballed platform (108) according to an angle rate command ?cmd (522), wherein the angle rate command ?cmd (522) is the acquisition phase angle rate command ?cmd—Acq (530) during the initial control period T and the tracking phase gimbal angle rate command ?cmd—Trk (532) after the initial control period T.

Abstract: A method and apparatus for controlling a gimbaled platform. The method comprises the steps of computing an acquisition phase gimbal angle rate command ?cmd—Acq from a measured LOS angle error ??LOS for an initial control period T while computing an estimated LOS angle rate {circumflex over (?)}LOS, computing a tracking phase gimbal angle rate command ?cmd—Trk using a controller having an output initialized with the estimated LOS angle rate {circumflex over (?)}LOS, and commanding the gimballed platform according to an angle rate command ?cmd, wherein the angle rate command ?cmd is the acquisition phase angle rate command ?cmd—Acq during the initial control period T and the tracking phase gimbal angle rate command ?cmd—Trk after the initial control period T.