Abstract: A Stationkeeping Mode for a spacecraft having a gimballed momentum wheel for attitude control in which the stationkeeping is divided into short duration open-loop stationkeeping maneuvers to correct orbital errors, and interlaced with closed-loop momentum management operations to manage momentum stored in the momentum wheel. The sequence of open-loop pulses are spaced to deadbeat a given dominant flexible motion of the spacecraft. Between and during the thruster pulses, the gimbaled momentum wheel is maintained in the closed-loop feedback control using gyro sensor or other inertial reference measurements. Limitations on the range of the momentum storage device and its torque capability are minimized by combining the maneuver thruster pulses with pulses from other thrusters, or by reducing the fire time of one of the maneuver thrusters, to minimize momentum buildup.

Abstract: A feedforward disturbance compensation system and method for use with a spacecraft that compensates for predictable disturbances exerted on the spacecraft. The invention comprises a model of the disturbance that is to be compensated, means for knowing when the disturbance will occur, and means for applying the feedforward compensation to the spacecraft in accordance with the model (a compensation profile). The present invention reduces pointing transients through the application of open-loop, feedforward compensation of actuators used to point the spacecraft. The shape of the feedforward profile is determined by an in-orbit test or is predicted by analytical means. For the in-orbit test case that compensates for an eclipse, for example, the spacecraft is flown through the eclipse once to determine the size and shape of the thermal shock disturbance and this information is stored in a spacecraft control processor.

Abstract: A method and system (12) for controlling the attitude of a spacecraft during its transfer orbit using an on-board, stand-alone, three-axes attitude determination and control system. The system utilizes a set of on-board sensors to define two independent angular measurements, which will initially identify the z-axis orientation of the spacecraft from an arbitrary attitude after launch vehicle separation. A set of three-axis gyros are then bias calibrated in order to measure the transverse rates of the spacecraft. The three-axis attitude of the spacecraft is continously determined by integrating the gyro outputs even if the Earth or Sun is not visible by an on-board sensor. A state estimator model is provided to determine the three-axis attitude of the spacecraft in the presence of large wobble and nutation. The system also utilizes a linear combination of the estimated attitude, rate and acceleration states to generate commanded rate increments with a pulse-width frequency modulator.

Abstract: A system located on board a geosynchronous satellite for detecting spin axis precession and developing thruster control signals for restoring the spin axis to the desired attitude. Fuel economy is enhanced by performing the major attitude correction at a time when the fuel expended contributes to that required for orbit inclination control.

Abstract: A pointing apparatus for a dual-spin spacecraft utilizing a first sensor for sensing the time of arrival of an inertial attitude reference as the spinning portion of the spacecraft rotates, and a second sensor for sensing the time of arrival of an index reference which relates the position of the despun portion with the spinning portion. A digital processor estimates the spin rate and phase of the spinning portion from the inertial attitude reference time of arrival, estimates the relative spin rate and phase between the spinning portion and the despun portion from the index reference time of arrival, and estimates the bearing friction bias torque on the motor means which controls the pointing direction of the despun portion of the spacecraft.