Aeronautics and astronautics – Spacecraft – Attitude control
Reexamination Certificate
2000-08-14
2002-11-19
Eldred, J. Woodrow (Department: 3644)
Aeronautics and astronautics
Spacecraft
Attitude control
C244S164000, C244S173300
Reexamination Certificate
active
06481671
ABSTRACT:
FIELD OF THE INVENTION
The invention relates generally to space vehicles and, more particularly, to structures and techniques for maintaining a space vehicle in a desired orientation during propagation in space.
BACKGROUND OF THE INVENTION
A spacecraft orbiting about the Earth, sun, or other body, or on an interplanetary mission, must typically maintain a predetermined attitude (i.e., physical orientation) to perform the function for which the spacecraft was designed. For example, the spacecraft may need to maintain a particular attitude so that: (a) an antenna on the spacecraft is pointing toward a desired region, (b) observation instruments on the spacecraft are directed toward an object being observed, and/or (c) solar panels on the spacecraft receive radiation from the sun. However, a number of forces normally act upon a spacecraft that can, if ignored, alter the attitude of the spacecraft from its desired orientation. These forces include, for example, the force of gravity, forces generated by the earth's magnetic field, and solar pressures generated by the sun's radiation impinging upon the surfaces of the spacecraft. These forces can generate undesired torques on the spacecraft that tend to rotate the spacecraft from its desired orientation. To deal with such forces, attitude control systems have traditionally been implemented.
Attitude control systems commonly rely upon angular momentum storage devices, such as control moment gyroscopes and/or reaction wheels, to track and compensate for undesired torques acting upon a spacecraft. Such devices react to external torques by storing momentum in an amount that is directly proportional to the level of external torque that is being applied to the spacecraft. A reaction wheel, for example, is a device that counteracts an external torque being applied to a spacecraft by changing the rate of rotation of the reaction wheel about its axis. The additional torque generated by accelerating the reaction wheel counteracts the undesired external torque acting upon the spacecraft in a manner that maintains the desired spacecraft attitude. A control moment gyro also counteracts external torques by storing angular momentum. However, the control moment gyro stores momentum by tipping the spin axis of a rotating member rather than accelerating a wheel. In addition to attitude maintenance functions, angular momentum storage devices can also be used to temporarily change the attitude of a spacecraft during special maneuvers or operations.
Over time, the amount of angular momentum stored within an angular momentum storage device can approach an upper limit. Once this limit is reached, it is impossible for the device to appropriately react to new external torques being applied to the spacecraft. Thus, “momentum dumping” is periodically employed to reduce the amount of momentum stored within the angular momentum storage device. Typically, momentum dumping is performed using thrusters on the spacecraft to generate torques that, when compensated for by the momentum storage device(s), eliminate the stored momentum within the device(s). The use of thrusters to perform momentum dumping, however, consumes a relatively large amount of fuel in the spacecraft. Consequently, additional fuel must be stored on the spacecraft to perform the momentum dumping. In addition, the use of thrusters for momentum dumping typically generates orbit variations for the spacecraft that necessitate additional station keeping activity, thereby consuming additional fuel.
Therefore, there is a need for a method and apparatus that is capable of efficiently maintaining a desired attitude for a spacecraft. There is also need for a method and apparatus that is capable of providing momentum management within a spacecraft in an efficient manner. Preferably, the method and apparatus will reduce the reliance on thrusters to perform attitude control functions on the spacecraft.
SUMMARY OF THE INVENTION
The present invention relates to a sunshield subsystem or assembly for use on a spacecraft that is capable of providing thermal control for the spacecraft while simultaneously performing attitude control functions for the spacecraft. Use of the sunshield subsystem can significantly reduce or eliminate the need to employ thrusters for momentum control in the spacecraft. In one embodiment, the sunshield subsystem reduces thruster propellant usage for momentum dumping by an order of magnitude, thus reducing the amount of thruster propellant that must be stored on the spacecraft to a small fraction of the fuel needed for station keeping. In a preferred approach, the sunshield subsystem is implemented on a spacecraft that needs a sunshield to provide thermal control for an application being performed by the spacecraft (e.g., to protect optics on-board the spacecraft). However, the sunshield subsystem can also be implemented on spacecraft that do not require additional thermal control, especially if the particular spacecraft design is known to nominally generate significant solar torques. The inventive principles are most advantageously used by spacecraft that utilize an inertial reference (e.g., the sun or stars) as opposed to an earth reference.
The sunshield subsystem includes a pair of shield members that are capable of being moved with respect to one another while the spacecraft is propagating through space. This movement capability is used to balance the solar torques acting upon the spacecraft, thereby significantly reducing the amount of torque that needs to be compensated for by momentum storage devices within the spacecraft. In one embodiment, the shield members are nominally flat structures comprising a thin film material supported by a rigid frame structure. At least one of, and preferably both of, the shield members are mounted on a moveable joint that allows each member, either separately or together, to pivot about an axis of rotation, such as relative to a pitch direction or plane. A motor unit is provided that allows one or both of the members to be controllably positioned about the corresponding axis. The relative positions of the shield members are adjusted in space so that an appropriate dihedral angle exists between the members to achieve a neutral stability condition for the spacecraft. Typically, an initial adjustment of the dihedral angle will be made when the spacecraft is first deployed to compensate for any inaccuracies in the initial shield member positions. Then, periodic adjustments are made to the dihedral angle during the operational life of the spacecraft to compensate for, for example, changes in the reflectivity of the sunshield material over time or damage to the shield members.
In one aspect of the invention, a sunshield subsystem is provided that is capable of balancing solar torques on a spacecraft in three independent rotational planes or directions (i.e., pitch, yaw, and roll). To achieve this, the shield members are made highly maneuverable so that they can be adjusted into many different mechanical configurations. For example, in addition to the dihedral angle adjustment capability that is used to balance torque in the pitch direction, a twisting capability is provided in the shield members for use in balancing torques in the yaw direction and a “rooftop” capability is provided for use in balancing torques in the roll direction. A controller is also provided for measuring a current solar torque associated with the spacecraft and for appropriately adjusting the shapes of the shield members to compensate therefor.
In another aspect of the invention, the maneuverability of the shield members is used to perform momentum dumping. That is, the shape of the shield members is adjusted in a manner that will generate a solar torque on the spacecraft that will reduce the amount of momentum stored within the momentum storage devices on-board the spacecraft. In this manner, momentum dumping is achieved without expending any thruster propellant. This momentum dumping technique can be used by itself or a hybrid system can be implemented that uses both su
Ball Aerospace & Technologies Corp.
Eldred J. Woodrow
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