Space Systems/loral, LLC

Philip Conway Hirschberg

Kam K Chan

Xenophon H Price

John James Rodden

Homer Darling Stevens

Mohammad Saghir Munir

John Hutton Cooper

Bruce Brumfield

George E White

David P Hong

Techniques for performing an orbital maneuver on a spacecraft by firing a thruster include executing a first double pulse, double coast firing (DPDC) sequence with the thruster prior to the orbital maneuver, firing the thruster for a duration of the orbital maneuver, and executing a second DPDC firing sequence with the thruster subsequent to the orbital maneuver.

Pre and post orbit maneuver pulses to reduce flexural oscillations

Stored momentum on a spacecraft is managed by determining a target profile of stored momentum as a function of time for the spacecraft; measuring a difference between a momentum value actually stored on the spacecraft and a desired momentum value, where the desired momentum value substantially conforms to the target profile at a particular time; reducing the difference by producing a torque on the spacecraft, where the torque results from selectively controlling at least one solar array position offset angle, the offset angle being an offset of at least one solar array of the spacecraft from a nominal sun pointing direction.

Spacecraft momentum management using solar array

Methods and apparatus for a spacecraft () orbiting about a celestial body such as the Earth to reacquire operational three-axis orientation with respect to that body. A method embodiment of the invention comprises determining () a set of actual conditions of the spacecraft, comprising a position of the spacecraft () in inertial space as a function of time and a set of angular rotation rates of the spacecraft () with respect to a coordinate frame of the spacecraft (), determining () an actual instantaneous direction of the sun with respect to the coordinate frame, and propagating () an estimated actual sun direction with respect to the coordinate frame as a function of time; determining () a desired sun direction with respect to the coordinate frame as a function of time; rotating () the spacecraft () and adjusting angular rotation rates of the spacecraft () so that an actual angle between the spacecraft () coordinate frame and the sun as a function of time substantially coincides with the desired sun direction with respect to the coordinate frame as a function of time; and, rotating () the spacecraft () around an axis defined by a line between the sun and the spacecraft until () the celestial body is observed by a celestial body sensor of the spacecraft.

Spacecraft three-axis attitude acquisition from sun direction measurement

An attitude control system comprising a controller and a noise screen device coupled to the controller. The controller is adapted to control an attitude of a vehicle carrying an actuator system that is adapted to pulse in metered bursts in order to generate a control torque to control the attitude of the vehicle in response to a control pulse. The noise screen device is adapted to generate a noise screen signal in response to the control pulse that is generated when an input attitude error signal exceeds a predetermined deadband attitude level. The noise screen signal comprises a decaying offset signal that when combined with the attitude error input signal results in a net attitude error input signal away from the predetermined deadband level to reduce further control pulse generation.

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Fremont, CA, United States of America

Philip Conway Hirschberg