Abstract: Methods and structures are provided that enhance the accuracy of the service attitude of an inclined-orbit spacecraft and, thereby, facilitate reduction of service error between a communication service area and the spacecraft's payload beam. The enhancement is realized by configuring a beacon-receiving antenna to have a beacon-receiving field-of-view that substantially matches a beacon-station window. Preferably, the beacon-receiving field-of-view is elongated and tilted to enhance its match with the beacon-station window in both size and orientation. The goals are also realized by configuring the beacon-receiving antenna to have a beacon-receiving field-of-view that is substantially smaller than the beacon-station window and successively steering a beacon-receiving boresight to successive beacon-receiving attitudes that maintain the beacon station within the beacon-receiving field-of-view over each solar day.

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: 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 spacecraft with a reaction wheel system can be autonomously safed by setting the solar wings to continuous tracking, determining a slew rate vector based on the total angular momentum, and slewing the spacecraft using the slew rate vector until commanded to stop autonomous safing. When the total angular momentum of the spacecraft to large to be handled by rotisserie, then the spacecraft is reoriented to align a suitable rotation vector with the system momentum. In a typical application, the spacecraft has a reaction wheel assembly with four wheels arranged to form a right regular pyramid. Two reaction wheels on opposite edges of the pyramid form a first pair and the two remaining reaction wheels forming a second pair.

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 satellite system includes a solar wing moveably connected to a satellite central body. A sensor, also coupled to the satellite central body, detects the movement of the body and generates a rate signal based on that movement. Additionally, an actuator, which controls momentum, is coupled to the satellite central body with maximum torque along the thermal shock axis. Subsequently, a rate-dominated thermal shock suppression controller, which is coupled to the satellite central body, receives the rate signal from the sensor to control the actuator.

Abstract: A spacecraft with a reaction wheel system can be autonomously safed by setting the solar wings to continuous tracking, determining a slew rate vector based on the total angular momentum, and slewing the spacecraft using the slew rate vector until commanded to stop autonomous safing. When the total angular momentum of the spacecraft to large to be handled by rotisserie, then the spacecraft is reoriented to align a suitable rotation vector with the system momentum. In a typical application, the spacecraft has a reaction wheel assembly with four wheels arranged to form a right regular pyramid. Two reaction wheels on opposite edges of the pyramid form a first pair and the two remaining reaction wheels forming a second pair.