Data processing: vehicles – navigation – and relative location – Vehicle control – guidance – operation – or indication – Aeronautical vehicle
Reexamination Certificate
2000-01-05
2002-02-12
Zanelli, Michael J. (Department: 3661)
Data processing: vehicles, navigation, and relative location
Vehicle control, guidance, operation, or indication
Aeronautical vehicle
C701S226000, C244S164000
Reexamination Certificate
active
06347262
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to systems and methods for controlling spacecraft, and in particular to a system and method for controlling the attitude and nutation of a spinning spacecraft using body stabilized spacecraft instruments.
2. Description of the Related Art
Satellites are in widespread use. These satellites include communications satellites used to deliver television and communications signals around the earth for public, private, and military uses, and surveillance satellites for collecting earth surface data.
In most cases, the location and orientation of each satellite must be controlled to perform its mission. Various satellite control system designs have evolved to achieve this purpose, including spin-stabilized and body stabilized designs. In spin-stabilized spacecraft designs, the spacecraft is stabilized by spinning satellite itself (or a momentum or reaction wheel within the satellite) about an axis. In body stabilized spacecraft designs, the spacecraft is stabilized in inertial space, typically with the use of an inertial guidance system which can include body mounted angular rate sensors to measure attitude rates and one or more staring inertial position sensors to measure instantaneous attitude (e.g. a sun sensor, star sensor, or earth sensor).
Many spacecraft provide dual-mode operation. Such satellites can be spin-stabilized to control attitude during high-thrust ascent maneuvers and in safe-hold modes, and body stabilized when performing their mission in operational orbit. To control the attitude of a spin-stabilized spacecraft having attitude control thrusters, the most efficient control of attitude and nutation is accomplished when synchronous (spin synchronous or nutation synchronous) pulsing is applied, as contrasted to continuous rate and/or position feedback control as is typically used on a body stabilized spacecraft.
Unfortunately, current dual-mode satellites use complex and expensive “spin unique” sensors and logic to stabilize the spacecraft when in the spin stabilized mode, and the inertial guidance system sensors and logic when in the body-stabilized mode. This increases the complexity, cost, and weight of the spacecraft.
What is needed is a simple method for autonomous attitude and nutation control of a spinning spacecraft that uses the simple continuous sensing and control law feedback compensation that is typically used on body-stabilized spacecraft. The present invention satisfies that need.
SUMMARY OF THE INVENTION
To address the requirements described above, the present invention discloses a method, apparatus, and article of manufacture for controlling the attitude and nutation of a spacecraft spinning about a first axis.
The method comprises the steps of measuring a spacecraft attitude rate &ohgr;
1
(t) about a second axis substantially orthogonal to the first axis; generating a rate correction value &agr;&ohgr;
1
(t)/&lgr;
0
, wherein &lgr;
0
is an inertial nutation frequency of the spinning spacecraft, and &agr; is a proportionality factor; determining a spacecraft attitude angle &phgr;
2
(t) about a third axis substantially orthogonal to the first axis and the second axis; generating an error signal e(t) as &phgr;
2
(t)+&agr;&ohgr;
1
(t)/&lgr;
0
; and applying a torque to rotate the spacecraft along the second axis. The article of manufacture comprises a program storage device implementing the instructions described above.
The apparatus comprises an attitude rate sensor for measuring the spacecraft attitude rate &ohgr;
1
(t) about a second axis substantially orthogonal to the first axis; means for measuring a spacecraft attitude angle &phgr;
2
(t) about a third axis substantially orthogonal to the first axis; a processor, in communication with the attitude rate sensor, for generating an error signal e(t)=&phgr;
2
(t)+&agr;&ohgr;
1
(t)/&lgr;
0
, wherein &lgr;
0
is an inertial nutation frequency of the spinning spacecraft, and a is a proportionality factor; and a torquer for applying a torque to the spacecraft along the second axis in accordance with the generated error signal e(t).
The foregoing not only provides the more efficient spinning control method, but also uses the same feedback structure and sensing instruments employed for body-stabilized control. Hence, a single sensor set and control structure may be used for both spinning and body stabilized modes of operation.
REFERENCES:
patent: 4758957 (1988-07-01), Hubert et al.
patent: 4931942 (1990-06-01), Garg et al.
patent: 5441222 (1995-08-01), Rosen
patent: 5597143 (1997-01-01), Surauer et al.
patent: 5794892 (1998-08-01), Salvatore
patent: 5826828 (1998-10-01), Fowell et al.
patent: 5826829 (1998-10-01), Holmes
Saeed Salma
Smay John W.
Yocum John F.
Gates & Cooper LLP
Hughes Electronics Corporation
Zanelli Michael J.
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