Abstract: An attitude determination system is provided for determining attitude values of a yaw-steering spacecraft, and includes a rate sensor, at least one attitude sensor, and a processor operable to receive measured spacecraft body rates from the rate sensor, to receive measured spacecraft attitude values from the at least one attitude sensor, and to calculate estimated spacecraft attitude values based on the measured spacecraft body rates and the measured spacecraft attitude values by using a Kalman filter that includes a plurality of attitude estimate error states, a plurality of gyro bias states, and a plurality of commanded rate dependent gyro error states.

Abstract: A system for providing precision thrust and sun tracking attitude control is provided. The system determines a proximity region and alternately engages either an ideal operational mode or a predictive operational mode based on whether a thrust trajectory vector is in the proximity region in order to provide attitude control. The proximity region is determined based on an angle between the thrust trajectory vector and a sun vector. For example, the angle is about 20-30 degrees. The system engages the predictive operational mode when the thrust trajectory vector enters the proximity region. When in the predictive operational mode, the system periodically re-calculates the thrust trajectory vector and determines where the thrust trajectory vector will exit the proximity region.

Abstract: Momentum control is maintained in a geosynchronous orbiting spacecraft that uses a plurality of reaction wheel assemblies and a plurality of magnetic torquers to control the spacecraft momentum, each orbit of the spacecraft being comprised of a set of time steps, by determining a current momentum error for a current time step of a current orbit by adding a system momentum change determined for an immediately preceding orbit to an average system momentum determined for the immediately preceding orbit, and then subtracting a magnetic control torque momentum change determined for the immediately preceding orbit, determining a current duty cycle for each of the magnetic torquers based on the current momentum error and on a torque value applied by each magnetic torquer at each time step of the immediately preceding orbit, and commanding each magnetic torquer to operate at the current time step in accordance with its respective determined current duty cycle, wherein the magnetic torquers apply a magnetic momentum c

Abstract: Method of and system for on-board substantially autonomous control for transferring a spacecraft from an initial orbit to a final geosynchronous orbit, by a trajectory that minimizes remaining transfer time and orbit transfer fuel. The spacecraft determines its orbit using a GPS-based system to determine the spacecraft orbital elements. Based on the measured orbit error, corrected co-state parameters are calculated and used to generate an updated thrust trajectory. The corrections are calculated using an innovative numerical procedure, carried out repetitively at a fixed interval until the target geosynchronous orbit is achieved.

Abstract: A spacecraft attitude control system and method in which an attitude controller is configured for sensing three-dimensional attitude of the spacecraft, and producing torque signals for stabilizing the spacecraft in a prescribed attitude in space. At least four mutually skew reaction wheels are rotated in response to the torque control signals for storing three-dimensional angular momentum, and speeds of rotation of the wheels are measured. A reaction wheel speed control processor, responsive to reaction wheel torque and speed for producing reaction wheel spin control signals, implements an infinity-norm algorithm for causing the nullspace components of wheel speed to re-distribute a desired three-dimensional stabilizing momentum of the spacecraft in such a manner that the maximum speeds of rotation of all the reaction wheels is minimized. As a result, periods between successive momentum dumping maneuvers are prolonged to the maximum possible.

Abstract: A spacecraft comprises a main body; at least one elongated solar wing extending from the main body, defining a generally flat plane and comprising a pair of longitudinally extending side edges; at least one other component or structure extending from the main body and spaced at a separation distance from the at least one solar wing; and at least one solar trim tab coupled to the at least one solar wing, linearly elongated and extending in the generally flat plane in a direction transversely away from one of the longitudinally extending side edges, and sized and positioned along the longitudinally extending side edge for counteracting or compensating for one or more types of disturbance torques.

Abstract: One embodiment of the present invention relates to a system for correcting spacecraft thermal distortion pointing errors. The system comprises one or more spacecraft sensors located at positions on a spacecraft and which are adapted to measure spacecraft parameters at those positions. The system also includes a spacecraft distortion prediction module, which is adapted to generate expected spacecraft thermal distortion parameter values and expected antenna thermal distortion pointing errors. Further, the system includes a spacecraft parameter processing module adapted to generate measured spacecraft thermal distortion parameter values from the measured spacecraft parameters, and an antenna pointing error calculation module adapted to calculate antenna pointing error correction commands. Finally, the system includes an antenna pointing control module adapted to receive the antenna pointing correction commands and control the adjustment of the antenna pointing using the correction commands.

Abstract: A velocity change (?V) thruster is operated on a spacecraft, which unavoidably causes attitude error. A reaction wheel (RWA) corrects the attitude. At the beginning of the thruster maneuver, the total attitude control momentum required to at least correct for the ?V thruster attitude errors over the duration of the entire maneuver is determined, and the RWA momentum may also be determined. Attitude control thrusters (REAs) are operated. The REAs are operated to correct at least the net ?V thruster induced attitude error, and preferably also to reset the RWA to its nominal momentum. The maneuver may be stationkeeping.

Abstract: A method for increasing the roll offset operating range for a spacecraft using an earth sensor operating in single scan mode includes the steps of moving the spacecraft to a first roll position, which has a roll angle that will cause the earth sensor to have a desired standard chord, switching the earth sensor to single scan mode by deselecting one of the earth sensor scans, whereby switching the earth sensor to a single scan mode, the earth sensor standard chord is locked at or near the desired standard chord. After the desired standard chord is set, the spacecraft is moved to a second roll position, which is a desired roll offset operating position for the spacecraft, and the earth sensor roll output (generated by using the single scan mode) is used to calculate the spacecraft roll at the roll offset operating position.

Abstract: The present invention relates generally to systems and methods for transferring a spacecraft from a first orbit to a second orbit. In accordance with one embodiment of the invention, the method comprises calculating thruster-off regions within an orbit transfer in which it is efficient to turn-off spacecraft thrusters, and in those thruster-off regions, turning off the spacecraft thrusters.

Abstract: One embodiment of the present invention relates to a method for increasing the roll offset operating range for a spacecraft using an earth sensor operating in a single scan mode. In accordance with this embodiment, the method comprises moving the spacecraft to a first roll position, which has a roll angle that will cause the earth sensor to have a desired standard chord. Next, the earth sensor is switched to single scan mode by deselecting one of the earth sensor scans. By switching the earth sensor to a single scan mode, the earth sensor standard chord is locked at or near the desired standard chord. After the desired standard chord is set, the spacecraft is moved to a second roll position, which is a desired roll offset operating position for the spacecraft. Finally, the earth sensor roll output (generating using the single scan mode) is used to calculate the spacecraft roll at the roll offset operating position.

Abstract: A velocity change (&Dgr;V) thruster is operated on a spacecraft, which unavoidably causes attitude error. A reaction wheel (RWA) corrects the attitude. At the beginning of the thruster maneuver, the total attitude control momentum required to at least correct for the &Dgr;V thruster attitude errors over the duration of the entire maneuver is determined, and the RWA momentum may also be determined. Attitude control thrusters (REAs) are operated. The REAs are operated to correct at least the net &Dgr;V thruster induced attitude error, and preferably also to reset the RWA to its nominal momentum. The maneuver may be stationkeeping.

Abstract: A method for maintaining three axis control of a geosynchronous spacecraft without body rate measurements using reaction wheel assemblies. Earth sensor assembly angle measurements are utilized for high-bandwidth roll and pitch control. A positive pitch momentum bias is stored in the reaction wheel assemblies. A gyroscopic feedforward torque is applied to rotate reaction wheel assembly momentum in a yaw/roll plane of the spacecraft at orbit rate. A dynamic mode that couples yaw and roll axes and that results from applying the gyroscopic feedforward torque and the high-bandwidth roll control is damped based on earth sensor assembly roll measurements.

Abstract: A method for controlling spacecraft velocity. A thruster torque gimbal angle command is calculated to provide a desired thruster torque demand command including a desired momentum adjust torque and a desired attitude control torque. A velocity change error between a commanded velocity change and an actual velocity change, a commanded force vector based on the error to cause the actual velocity change to approach a commanded velocity change, and force gimbal angle commands and force thrust commands to achieve the commanded force vector are calculated; and/or an error between gimbal angles and thrust levels and reference gimbal angles and thrust levels, reference gimbal angle commands and reference thrust commands to drive actual gimbal angles and thrust levels toward reference gimbal angles and reference thrust levels are calculated. A total gimbal angle command is calculated by adding the torque gimbal command with at least one of the force gimbal angle command and the reference gimbal angle command.

Abstract: A spacecraft (10) carries a solar panel (17) which rotates to follow the sun, and also carries various thrusters (20). Thruster plume impingement on the solar panel affects the torque applied to the spacecraft body (12) in a manner which depends upon solar panel angle. The errors in the thrust during stationkeeping tend to perturb attitude, especially early in the maneuver, because of the delay inherent in the attitude control loop. A torque bias is summed with the residual torque demand signal to correct for the errors in torque. The torque bias signal is generated by a Fourier model of the torques, updated by an adaptive tuning filter, so that successive stationkeeping maneuvers progressively adapt the amplitude and phase of the Fourier coefficients in a manner which tends to minimize the residual torque demand and attitude error. Thus, the torque bias signal automatically approaches the correct value.

Abstract: A spacecraft includes a three-axis attitude control system. When velocity change thrusters are fired, their plumes impinge on a solar array, at angles which vary with the solar array position. This causes disturbance torques which vary with the solar array position. Disturbance torque information signals or torque bias signals which depend upon the solar array angle are summed with the torque demand signals which control the attitude control system during firing of the velocity change thrusters, to modify the attitude correction torques. The bias torque signals are generated by a Fourier processor based upon stored Fourier coefficients together with signals from a solar array angular position sensor.

Abstract: A three-axis stabilized spacecraft includes a plurality of primary attitude control thrusters, the torque vectors of which lie in, or parallel to a primary plane. It also includes at least two more secondary attitude control thrusters, the torque vectors of which lie in a secondary plane which is not parallel to the primary plane. The control system produces attitude error signals, which are processed with a PID characteristic to produce impulse demand signals, all in known fashion. The impulse demand signals are transformed into an auxiliary coordinate system, in which two of the three auxiliary axes lie in the primary plane, and the third is orthogonal thereto. One of the secondary thrusters is selected, which has, along the third auxiliary axis, the largest torque magnitude and the same sign as the transformed impulse demand.