Abstract: Methods and apparatus for a spacecraft (1) orbiting about a celestial body such as the Earth to reacquire operational three-axis orientation with respect to that body.

Abstract: A momentum wheel configuration that may be used on a spacecraft having a momentum bias that prevents zero wheel speeds. The momentum wheel configuration comprises three or more momentum wheels. The momentum wheels are oriented such that their wheel axes are not parallel to each other. The momentum wheels are oriented such that their axial momentum vectors are not coplanar. None of the momentum wheels have their wheel axis parallel to the momentum bias direction of the spacecraft. No two of the momentum wheels are oriented such that their axial momentum vectors are coplanar with the momentum bias direction of the spacecraft.

Abstract: A momentum wheel configuration that may be used on a spacecraft having a momentum bias that prevents zero wheel speeds. The momentum wheel configuration comprises three or more momentum wheels. The momentum wheels are oriented such that their wheel axes are not parallel to each other. The momentum wheels are oriented such that their axial momentum vectors are not coplanar. None of the momentum wheels have their wheel axis parallel to the momentum bias direction of the spacecraft. No two of the momentum wheels are oriented such that their axial momentum vectors are coplanar with the momentum bias direction of the spacecraft.

Abstract: A system and method for RF autotracking multiple antennas (preferably located on a spacecraft) to compensate for disturbances experienced by the antennas. The system and method estimate high frequency errors associated with all of the antennas. The high frequency errors may be estimated using a currently selected antenna and a high-pass filter. The high frequency errors may be estimated using a sensor mounted on the spacecraft, such as a gyro or star tracker, or may be estimated using information, such as planned thruster firings, for example, from a spacecraft attitude control system. Alternatively, the high frequency information may be estimated using any combination of data from these sources. Low frequency errors are estimated using measurements from each selected antenna. The algorithm implemented in the present invention explicitly accounts for the frequency content of each disturbance source.

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: Systems and methods that limit the effects of actuator saturation to certain body axes of a spacecraft. A plurality of actuators, or momentum wheels, disposed on the spacecraft are oriented non-parallel to x, y, and z axes of the spacecraft. The plurality of actuators each produce torque that is applied to the spacecraft. Outputs of the plurality of actuators or momentum wheels are processed to control the respective actuators or momentum wheels to manage actuator saturation so that saturation does not produce attitude errors about one or more predetermined axes of the spacecraft.

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.