Hughes Aircraft Company

Hughes Electronics Corporation

Douglas J Bender

John F Yocum, Jr

Stuart F Bockman

Piyush R Shah

Richard A Fowell

Bruce N Eyerly

John Joseph Anagnost

David S Uetrecht

Loren I Slafer

Dan Y Liu

William L Burkett, Jr

William F Hummel

Thomas F Brozenec

James D Brehove

Thomas R Parks

William A Nakano

Kenneth R Obert

Tiffany Bowles

A sun-referenced safe-hold control system for momentum-biased satellites employs a safe-hold processor responsive to a set of sun sensors to detect attitude errors of the solar wings. Safe-hold processor provides independent control logic which rotates the solar wings to ensure that sunlight is maintained incident on the solar wings to maintain solar power in the safe-hold mode. In addition, the safe-hold processor provides independent control of the momentum wheel to place the spacecraft body in a passively stable nutational state. Passive nutational stability inherent in a momentum-biased satellite about two axes of the spacecraft enables the safe-hold processor to actively control the spacecraft about only one of its axes. The spacecraft safe-hold control system increases the period in which spacecraft health is guaranteed by adding active, closed-loop control of the solar wings.

Sun-referenced safe-hold control for momentum biased satellites

A system and method for acquiring Sun pointing in a three-axis stabilized spacecraft including slewing at least one solar wing until the Sun is detected, determining an initial Sun vector, performing one or more rotations of the spacecraft body so as to bring the instantaneous Sun vector coincident with a preferred final Sun vector, at least one rotation about an optimal axis, and slewing at least one solar wing to a preferred attitude relative to the Sun. In one embodiment, the optimal axis is chosen so as to minimize the time required to achieve an optimal thermal attitude. In another embodiment, the optimal axis is chosen so as to minimize the time required to align the instantaneous Sun vector with the final desired Sun vector. In a further embodiment, the optimal axis is chosen so as to maintain Earth lock.

Spacecraft acquisition of sun pointing

A system and method for slewing a spacecraft with a reaction wheel system including determining initial total momentum, determining maximum forward and reverse rates that can be realized within the reaction wheel system maximum storage capacity, determining a slew rate and slew direction from the maximum forward and reverse rates so as to minimize slew time for known slew angles or maximize slew rate for arbitrary slew angles, and slewing the spacecraft.

System and method for utilizing stored momentum to optimize spacecraft

A multi-loop control system for a gimballed antenna that employs devices for measuring both absolute line-of-sight (an autotrack receiver or beacon tracker) and relative angular position (a resolver). The control system uses both signals simultaneously, thereby increasing the performance and pointing accuracy capability. Two control loops operate simultaneously to provide for optimum performance. The first loop is an inner high-bandwidth control loop that uses the relative gimbal angle measurement to control pointing of the antenna along a precommanded profile. The inner loop may run alone to provide for coarse pointing. When available, the line-of sight measurement is used in a low-bandwidth outer loop to provide corrections to the command profile of the inner loop. Control logic is provided that allows switching between several control modes. By using the present invention, antenna tracking control performance is maximized, especially in the presence of attitude disturbances of a spacecraft or significant flexible interactions.

Gimbal control system

A method and system for determining the general three-axis attitude measurement for a spacecraft is disclosed. The system provides a cost-effective attitude measurement system for geosynchronous momentum bias spacecraft which need a continuously updated, moderately accurate, yaw measurement in addition to the usual roll and pitch measurements. The continuous yaw measurement enables control of the spacecraft yaw axis in the presence of disturbance torques. A sun sensor generates a signal which may be used to calibrate a continuously updated yaw measurement derived from the combination of signals generated by an Earth sensor and a space-to-ground link.

Spacecraft attitude determination using sun sensor, earth sensor, and

This invention discloses a method for compensating for the disturbance torque on an orbiting satellite (10) from the interference of the earth's magnetic field lines with particular current loops associated with the satellite (10). The current flow in these current loops on the satellite (10) is measured by a current sensing device (42). A scaling factor (44,46) is applied to the measured currents to derive an estimated disturbance torque in particular axes of the satellite's attitude. The estimated magnetic disturbance torques are applied, along with other control torques, to a satellite actuation device (38). The actuation device (38) actuates the satellite (10) in the particular axes to compensate for the disturbance torques being applied to the satellite (10).

Method and apparatus for compensating for magnetic disturbance torques

A method of attitude measurement for an artificial satellite (100) utilizes one or more star trackers (12) together with an earth sensor (30). Periodic updates of satellite orbital information, either propagated onboard or from a ground station are combined with earth and star position coordinate data to provide a continuous and accurate measurement of the spacecraft body 3-axis attitude. The method can be used for ground-based attitude determination or onboard closed loop control systems.

Method of attitude determination using earth and star sensors

This invention discloses a method for determining when an orbiting satellite (10) is eclipsed from the sun in order to remove control torques to the satellite (10) which compensate for the disturbance of solar pressure on the satellite (10). A current measuring device (46) measures the current traveling through a particular circuit associated with the satellite (10) which is indicative of the satellite batteries being discharged, as would occur during an eclipse. The measured current is applied to a threshold logic circuit (48) which sends a signal to a control compensator (36) if the measured current exceeds a predetermined threshold level. Consequently, the compensation provided by the control compensator (36) removes the compensation for compensating for solar pressure when the satellite (10) is in an eclipse. In a second implementation, the threshold logic circuit (48) is replaced with a proportionality logic circuit to compensate for the effects of partial eclipses.

Method and apparatus for compensating for solar torque transients on a

The three axes thruster modulation (8) of the present invention accepts three axes of input torque commands or angular acceleration commands and generates thruster selection and thruster timing (40) information which is used to fire thrusters (48) for the purpose of spacecraft attitude control and velocity change maneuvers. The modulation logic (8) works in all three axes simultaneously and is suitable for use with an arbitrary thruster configuration, including a configuration in which individual thrusters or thruster groups do not produce torques about mutually orthogonal axes. After thruster selection and on-times have been determined, the modulation logic (8) uses this information to compute a best estimate of the actual rate change (42) which is then compared to the commanded rate change (44) to develop a residual unfired rate change. The residual unfired rate change is retained and fed forward for addition to the subsequent acceleration command in computing the commanded rate change for the subsequent sample period (10, 20, 30). As a result, all commanded rate changes will ultimately be fired even if the commanded rate change is below a minimum rate change achievable.

Three axis thruster modulation

Method and apparatus for controlling the two-axis rate and position as well as the linear elevation of a momentum wheel platform supported by three jackscrews driven by stepper motors. By tilting the momentum wheel, a control torque is produced in the plane orthogonal to the nominal wheel spin axis. When all three actuators are operational, the control transformations are made unique by constraining the sum of the actuator extensions to be zero. When one actuator has failed, the above constraint is eliminated and the failed actuator is constrained to its measured value. Momentum wheel platform rotational rate commands generated by the roll/yaw control system are first integrated to give desired gimbal angles, and then transformed into jackscrew extensions. Jackscrew step commands are computed as the difference between these desired extensions and the current positions of the jackscrews. The position of each jackscrew is maintained by accumulating the step commands. The desired extensions are also back-transformed to compute gimbal angles that are fed back to the roll-yaw state control system. In ground commanded stepping, the momentum wheel platform steering method operates exactly as described above, except that relative jackscrew step commands from a ground station are utilized instead of the step commands computed in the rate following configuration.

Momentum wheel platform steering system and method

This invention discloses a method for maintaining a desirable orientation of the solar wings of an orbiting satellite (10) relative to the sun. A dual array of sun sensors (26,28) is positioned on a body (12) of the satellite (10) in order to get a measurement of the position of the sun relative to the body (12) once per every orbit of the satellite (10). In addition, an estimate of the position of the solar wings (16,17) relative to the body (12) of the satellite (10) is attained. The sun-to-body angle is then subtracted from the body-to-wing angle to drive an error signal which is applied to a wing driver mechanism in order to maintain the solar wings (16,17) of the satellite (10) in a proper orientation relative to the sun.

Method and apparatus for controlling a solar wing of a satellite using a

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. The spinning portion spin rate and phase estimates are added with the relative spin rate and phase estimates to produce an estimate of the despun portion spin rate and phase, and the despun portion spin rate and phase estimates and the friction bias torque estimates are subtracted from commanded despun portion spin rate, phase and friction bias torque states. A torque command is generated for controlling the motor means from the subtracted estimates.

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Redondo Beach, CA, United States of America

Douglas J Bender