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: 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: An attitude sensing system utilizing simplified techniques and apparatus includes a Kalman filter which receives signals from an inertial measurement unit, a GPS receiver, and an integrated optical assembly. The output vector of the filter includes estimates of attitude misalignments and estimates of gyro drifts corresponding to the axes of the inertial measurement unit. The optical assembly includes a sensor array providing signals to the filter representing detection of the Earth's horizon or the center of the Sun. More particularly, a local vertical vector, computed from fore and aft detections of the Earth's horizon, is used in combination with GPS received signals to initially determine attitude by means of gyrocompassing. This attitude information is thereafter maintained by the inertial measurement unit and azimuth error resulting from drift of the inertial measurement unit and the initial gyrocompassing error is corrected by detections of the Earth's horizon and the Sun.

Abstract: An inertial navigation system with automatic redundancy and dynamically calculated gyroscopic drift compensation utilizes three, two-degree of freedom gyroscopes arranged whereby any two of the gyroscopes form an orthogonal triad of measurement sensitive axes. The input axes of the three gyroscopes form three pairs of parallel input axes, each pair of parallel input axes corresponding to one axis of the orthogonal triad of axes. The three gyroscopes are operated in a plurality of preselected combinations of both clockwise and counter clockwise directions, thus changing the direction of the angular momentum vector by 180.degree.. Parity equations are formed from each pair of gyroscope outputs whose measurement sensitive axes are parallel. The parity equations include combinations of gyroscope pairs that have been operated in both the clockwise and counterclockwise directions.