Aeronautics and astronautics – Spacecraft – Attitude control
Reexamination Certificate
1999-06-24
2001-03-06
Barefoot, Galen L. (Department: 3644)
Aeronautics and astronautics
Spacecraft
Attitude control
Reexamination Certificate
active
06196502
ABSTRACT:
TECHNICAL FIELD
The present invention relates generally to positioning of a spacecraft, and more particularly relates to attitude control of a spinning spacecraft.
BACKGROUND ART
Attitude control of a spin stabilized spacecraft has traditionally been accomplished by pulsed thrusters which provide spin phased moments. This works well when the operational spin axis is normal to the orbit plane. The payloads in some spacecraft, however, require a nadier pointing spin axis. In order to alleviate the necessity for continuous thrusting to maintain the orbital angular rate, momentum canceling wheels are hardmounted to current spacecraft to create a zero momentum system. When it is desired to slew the spin axis from one pointing direction to another, reaction control thrusters are used to create the starting and stopping moments. If solar radiation pressure creates undesired external moments, these thrusters are also used to cancel such moments.
While reaction control thrusters are necessary, it would be desirable to use them as little as possible. There is only a limited amount of propellant available to fire the thrusters and the more that the thrusters must be fired, the less fuel is available for later maneuvers. Further, there are inherent uncertainties involving firing a thruster on time and sophisticated methods are required in order to compensate for these uncertainties.
U.S. Pat. No. 5,441,222 to Rosen discloses an attitude control system that is intended to minimize the use of thrusters. The attitude control system utilizes an agile gimbaled momentum wheel to provide attitude control moments and cancel the momentum of a spinning satellite and thus provide a zero overall system momentum. The disclosed momentum wheel assembly executes a series of coning motions to provide the correct control moments. The coning motion necessary for body attitude control and momentum control tends to require significantly more angular motion than that needed for wobble control. This is especially true when balancing techniques are employed (either passive or active) which can reduce the spacecraft wobble to an arbitrarily small value. With this configuration, the momentum wheel gimbal actuators typically need to provide large excursion, high rate, long life, and low disturbance torque position control. This results in a complex and costly implementation, such as disclosed in U.S. Pat. No. 5,820,078 to Harrell.
SUMMARY OF THE INVENTION
Accordingly, it is an object of the present invention to provide an attitude control system for a spinning spacecraft that requires smaller excursion motions.
It is a further object of the present invention to provide an attitude control system for a spinning spacecraft that will improve the durability of the assembly.
It is another object of the present invention to provide an attitude control system for a spinning spacecraft that has prolonged life and improved accuracy over prior assemblies.
In accordance with these and other objects of the present invention, an attitude control system for a spinning spacecraft is provided. The spinning spacecraft has a despun platform in rotational communication with the body of the spinning spacecraft. A momentum control system is mounted on the despun platform. The momentum control system includes a gimbal having first and second portions that are respectively pivotable about first and second orthogonal axes. The gimbal is in communication with a spinning momentum wheel. A first actuator is connected between the first portion of the gimbal and the despun platform for applying a first control moment to the despun platform about the first axis. A second actuator is connected between the second portion of the gimbal and the despun platform for applying a second control moment to the spacecraft about the second axis. The first and second actuators are respectively secured to the first and second portions of the gimbal at locations along the first and second axis respectively to avoid the undesirable application to the spacecraft of unwanted secondary moments. The first and second control moments are operative to control the attitude of the spacecraft.
These and other features and advantages of the present invention will become apparent from the following description of the invention when viewed in accordance with the accompanying drawings and appended claims.
REFERENCES:
patent: 4375878 (1983-03-01), Harvey et al.
patent: 4911385 (1990-03-01), Agrawal et al.
patent: 5112012 (1992-05-01), Yuan et al.
patent: 5441222 (1995-08-01), Rosen
patent: 5820078 (1998-10-01), Harrell
Barefoot Galen L.
Gudmestad Terje
Hughes Electronics Corporation
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