Satellite attitude determination and control system with agile beam sensing

• US 4,883,244 A
• Filed: 12/23/1987
• Issued: 11/28/1989
• Est. Priority Date: 12/23/1987
• Status: Expired due to Term

• Current Assignee: Hughes Electronics Corporation (Rtx Corporation)
• Original Assignee: Hughes Aircraft Company (Rtx Corporation)
• Inventors: Fowell, Richard A., Huang, Ken Y., von der Embse, U. A., Allen, Gene E., Mitchell, Mark P., Chang, Donald C. D., Challoner, A. Dorian, Hayden, Joseph H.
• Primary Examiner(s): NOT, DEFINED
• Assistant Examiner(s): Corl, Rodney
• Application Number:
  US07/137,225
• Time in Patent Office: 706 Days
• Field of Search: 244/164, 244/166, 244/171, 364/449, 364/455, 364/456, 364/459, 342/352-358
• US Class Current: 244/171
• CPC Class Codes: B64G 1/244 Spacecraft control systems
• CPC Class Codes: B64G 1/28 using inertia or gyro effect
• CPC Class Codes: B64G 1/281 Spin-stabilised spacecraft
• CPC Class Codes: B64G 1/32 using earth's magnetic field

Abstract

Agile (electronically steerable) beam sensing with associated on-board processing, previously used exclusively for positioning of antennas for beam formation and tracking in communications systems, is now also used for satellite active attitude determination and control. A spinning satellite (100) is nadir oriented and precessed at orbit rate using magnetic torquing determined through use of an on-board stored magnetic field model (520) and attitude and orbit estimates (212). A Kalman filter (211) predicts parameters (202, 203) associated with a received signal (204) impinging on the satellite'"'"'s wide angle beam antenna (201). The antenna system measures the error between the parameter predictions and observed values and sends appropriate error signals (207) to the Kalman filter for updating its estimation procedures. The Kalman filter additionally outputs the spacecraft attitude error signals (215) to an attitude control law (213), which determines commands to attitude-altering magnetic torque elements (220) to close the control loop via the spacecraft dynamics (230).

Claims

1. An on-board attitude determination system for a spacecraft orbiting a celestial body comprising:

means for estimating the orbit and attitude of the spacecraft relative to a plurality of remote sources of coherent energy signals;

an agile beam sensor sensing the locations of the remote sources;

means for pointing the agile beam sensor at the estimated relative locations of the remote sources;

means for determining error signals representative of the difference between the estimated locations of the remote sources and actual locations of the remote sources as sensed by the agile beam sensor; and

said estimating means being adapted to receive said error signals and update the predicted orbit and attitude of the spacecraft.

2. The system of claim 1 wherein said agile beam sensor comprises an electronically steerable phased array antenna.

3. The system of claim 2 wherein the plurality of remote sources include orbiting satellites.

4. The system of claim 2 wherein the sensed locations includes azimuth, elevation and range of each remote source relative to the phased array antenna.

5. The system of claim 2 wherein said estimating means includes a Kalman filter.

6. A method of attitude control for a spacecraft orbiting a celestial body comprising the steps of:

orienting the spacecraft such that it spins about a spin axis oriented substantially in the orbit plane and pointing substantially at nadir; and

precessing the spin axis at a rate equal to the rotational orbit rate such that the spin axis of the spacecraft will remain substantially fixed relative to a celestial body centered coordinate frame rotating at the rotational orbit rate throughout an entire orbit for the spacecraft about the celestial body, said precessing step including the steps of;

predicting values of preselected directional parameters relative to the spacecraft of coherent energy signals radiating from a plurality of known remote sources;

observing the energy signals, determining values of the directional parameters from the observed energy signals, calculating differences between observed and predicted values of the directional parameters, and generating error signals indicative of the differences;

updating the predicted parameters in accordance with the error signals and generating a spacecraft attitude state estimate from updated parameter values;

comparing the state estimate to a reference state and generating a plurality of command signals in accordance with the comparison results; and

generating torques upon the spacecraft such that the spin axis will precess about an axis substantially normal to an orbit plane, such as to maintain nadir pointing of the spacecraft.

7. The method of claim 6 wherein the observed values of preselected directional parameters are based on an electronically steerable agile beam of coherent energy.

8. An on-board control system for a spacecraft orbiting a celestial body comprising:

means for predicting values of preselected directional parameters of coherent energy signals radiated from a plurality of known remote sources;

sensing means coupled to the means for predicting and operative to observe the energy signals, to measure differences between observed and predicted values of the directional parameters, and to generate error signals indicative of the difference;

the means for predicting including means, coupled for receipt of the error signals, for updating the predicted parameters in accordance with the error signals and for generating a spacecraft orbit and attitude state estimate from updated parameter values; and

decision means, coupled for receipt of the state estimate, operative to compare the state estimate with a reference state and to generate at least one command signal in accordance with the comparison results.

9. The control system of claim 8 further comprising torque generating means operative to generate an attitude-adjusting torque on the spacecraft of a magnitude and direction determined by the at least one command signal.

10. The control system of claim 8 further comprising force generating means operative to generate an orbital velocity adjusting force on the spacecraft of a magnitude and direction determined by the at least one command signal.

11. The control system of claim 9 wherein the celestial body has a magnetic field and wherein the torque generating means comprises at least one magnetic moment generating means coupled for receipt of a respective command signal and operative to generate a moment having magnitude and direction appropriate for generating the torque in conjunction with the magnetic field of the celestial body.

12. The control system of claim 8 wherein the means for predicting includes a Kalman filter for estimating the directional parameters and for generating an updating the attitude state estimate.

13. The control system of claim 8 wherein the celestial body has a magnetic field and the decision means includes a stored model of the magnetic field of the celestial body.

14. The control system of claim 13 wherein the model is derived and continuously adjusted using agile-beam-sensed attitude and orbit state estimates.

15. The control system of claim 8 wherein the reference state is arranged such that a spin axis of the spacecraft will be maintained nadir pointing throughout each spacecraft orbit of the celestial body.

16. The control system of claim 11 wherein the magnetic moment generating means comprises a substantially cylindrical permeable control rod having a control winding thereon.

17. The control system of claim 16 wherein the control winding encircles an air impermeable core.

18. The control system of claim 16 wherein the control winding encircles a permeable magnetic core.

19. The control system of claim 16 wherein there are a plurality of control rods and each control rod is orthogonal to every other of the plurality of control rods.

20. The control system of claim 19 wherein there are three mutually orthogonal control rods.

21. The control system of claim 8 wherein the sensing means further comprises a phased array, steerable beam antenna.

22. The control system of claim 21 wherein the directional parameters include azimuth, elevation, and range of each remote source relative to the phased array antenna.

23. The control system of claim 22 wherein each remote source radiates coherent radio frequency information signals.

24. The control system of claim 11 wherein the decision means further comprises:

means for generating the reference state;

first, second and third comparison means, each having first and second inputs respectively coupled to the means for generating and the means for predicting, the first comparator having an output coupled to precession control means, the second comparator having an output coupled to spin rate control means and the third comparator having an output couplred to transverse rate control means; and

magnetic field modeling means having an input coupled to the means for predicting;

and wherein the magnetic-moment generating means comprises torque summing and commutation means for generating at least one command signal and having inputs respectively coupled to outputs of the precession control means, the spin rate control means, the transverse rate control means and the magnetic field modeling means, a fourth comparison means having a first input coupled to an output of the means for generating, a second input coupled to an output of the torque summing and commutation means, and transforming means for converting magnetic moment information present at an output of the fourth comparison means from despun coordinate form to spinning body coordinate form and having inputs respectively coupled to an output of the fourth comparison means to the means for predicting.

25. The control system of claim 8 wherein the plurality of known remote sources includes orbiting satellites.