Aeronautics and astronautics – Spacecraft – Spacecraft formation – orbit – or interplanetary path
Reexamination Certificate
1999-11-26
2001-11-06
Jordan, Charles T. (Department: 3644)
Aeronautics and astronautics
Spacecraft
Spacecraft formation, orbit, or interplanetary path
C244S171000, C244S173300, C244S164000
Reexamination Certificate
active
06311929
ABSTRACT:
BACKGROUND
The present invention relates generally to spacecraft attitude pointing methods, and more particularly, to spacecraft attitude pointing methods that provide for stepping multiple appendages to reduce spacecraft attitude pointing disturbances caused by appendage stepping and cancellation of solar array slew disturbances.
A class of spacecraft known as three-axis stabilized spacecraft employ a solar array to generate power for the spacecraft. The solar array must be maintained in a position normal to the sun to absorb the optimum amount of radiation. Because the solar array is maintained normal to the sun, a servo controlled stepping mechanism, such as a stepping motor and an appropriate gear train, is typically employed to cause the solar array to track the sun while the spacecraft is in constant rotation relative to the sun in an orbit about the earth. Other types of attitude control mechanisms, such as dc motors, prove to be relatively difficult to control and are heavy. However, in theory, servo controlled dc motors would not generate oscillation. It is desirable to use stepper motors because stepper motors are relatively simple to control, reliable, lightweight and well adapted to continuous use.
One of the major problems with the use of stepping motors is that the stepping action can excite a highly flexible array such that oscillation is induced within the spacecraft. The induced oscillation is particularly critical in spacecraft where absolute platform stability is desirable or required, such as platforms for high resolution optical imaging systems. Vibrations can cause deterioration of any inertia-sensitive operations of a spacecraft. Therefore, it is desirable to solve the problem of induced oscillation caused by a stepper motor.
U.S. Pat. No. 4,843,294 entitled “Solar Array Stepping to Minimize Array Excitation” assigned to the assignee of the present invention discloses one way to improve spacecraft attitude pointing. The method disclosed in U.S. Pat. No. 4,843,294 deadbeats individual appendage oscillations. As such, the stepping of the solar array wings were stepped in a manner that minimized their individual oscillations. The present invention improves upon the teachings of U.S. Pat. No. 4,843,294.
More particularly, and in accordance with the teachings of U.S. Pat. No. 4,843,294, mechanical oscillations of a mechanism containing a stepper motor, such as a solar-array powered spacecraft, are reduced and minimized by the execution of step movements in pairs of steps. The period between steps is equal to one-half of the period of torsional oscillation of the mechanism. Each pair of steps is repeated at needed intervals to maintain desired continuous movement of the portion of elements to be moved, such as the solar array of a spacecraft. In order to account for uncertainty as well as slow change in the period of torsional oscillation, a command unit may be provided for varying the interval between steps in a pair.
Furthermore, solar arrays are Sun tracking, while satellite payloads are Earth tracking. This means the solar arrays rotate with respect to the body of the spacecraft. Every step in the rotation causes a disturbance. As solar arrays become physically larger, so do the disturbances caused by rotation the solar arrays. Previous systems developed by the assignee of the present invention relied entirely on feedback to reduce the disturbances. The present invention takes apriori knowledge of an event (solar array step) and uses that knowledge to reduce the disturbance.
Accordingly, it is an objective of the present invention to provide for spacecraft attitude pointing methods that provide for stepping multiple appendages to reduce spacecraft attitude pointing disturbances caused by appendage stepping and cancellation of solar array slew disturbances.
SUMMARY OF THE INVENTION
To accomplish the above and other objectives, the present invention provides for a method of stepping multiple appendages (North and South solar array wings) that reduces spacecraft attitude pointing disturbances caused by appendage stepping and thus improves spacecraft attitude pointing. In contrast to the method disclosed in U.S. Pat. No. 4,843,294, the present invention uses multiple appendages to deadbeat the impact on the spacecraft body oscillation. That is, the system disclosed in U.S. Pat. No. 4,843,294 controls the time between steps of the same appendage, while the present invention controls the time between steps of different appendages. The present invention is advantageous when the same mechanism does not require two successive steps. The present invention provides for a simple method that improves spacecraft attitude pointing if multiple appendages are stepped and which can be optimized easily while in-orbit using ground commands.
The North and South solar array wings are stepped at a rate required for sun tracking. The period between the North and the South wing steps is chosen that deadbeats the flexible appendage disturbance imparted to the spacecraft body by the solar array wings. The deadbeat interval is less than the step rate required for sun tracking. Thus, spacecraft attitude pointing is improved without requiring changes to the sun tracking rate of each individual wing. Also, the timing between the first and second appendage steps may be ground commanded so that uncertainties in the flexible properties of the solar array wings can be optimized easily while the spacecraft is on-orbit.
Thus, the present invention phases the motion of the solar array wings so that the combined effect on the spacecraft body is minimized. Stepping of the motion of one wing is timed with respect to stepping of the second wing so that the spacecraft body oscillations are minimized. The phased motion of the North and South solar array wings combine to minimize the motion of the spacecraft body. Thus, both solar array wings may oscillate but the spacecraft body of the does not. This improves the pointing accuracy (antenna pointing) of the spacecraft.
The present invention also provides for a second method that cancels solar array slew disturbances exerted on the spacecraft body. The second method reduces the magnitude of the disturbance to the spacecraft body, which results in smaller pitch errors.
The second method involves predicting the disturbance torque exerted on the spacecraft due to the stepping of a flexible appendage, and having the spacecraft actuators compensate for the disturbance torque before it creates a pointing error. Since this torque is relatively small, current control systems allow the disturbance torque to produce a pointing error and then cancel the pointing error with feedback control. However this invention cancels the predictable disturbance torque before it produces a pointing error, improving the spacecraft pointing performance.
The term “feedback” means to calculate a control signal by processing sensor data. Conversely, the term “feedforward” means to calculate a control signal without using sensor data. For example, if a body was exposed to a predictable disturbance torque, feedforward control would apply and equal and opposite attitude control torque to cancel the disturbance torque before it created a pointing error. Feedforward control is not used alone, but rather is used with feedback control in order to make the feedback control more effective.
The present invention may be used with geosynchronous orbit spacecraft having large solar arrays wherein the solar arrays continuously track the sun. The present invention may also be used with apparatus where it is desirable to minimize vibration introduced by stepped excitation.
REFERENCES:
patent: 4834294 (1989-05-01), Bhat et al.
patent: 5540405 (1996-07-01), Bender et al.
patent: 5610848 (1997-03-01), Fowell
patent: 5687933 (1997-11-01), Goodzeit et al.
patent: 5697582 (1997-12-01), Surauer
patent: 5816540 (1998-10-01), Murphy et al.
Higham John
Kazimi Fatima
Leung Yat Fai
Liu Tung
Reckdahl Keith
Jordan Charles T.
Karambelas Anthony W.
Space Systems Loral, Inc.
Steele George L.
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