Attitude pointing error correction system and method for geosynchronous satellites

• US 4,911,385 A
• Filed: 04/30/1987
• Issued: 03/27/1990
• Est. Priority Date: 04/30/1987
• Status: Expired due to Fees

• Current Assignee: Intelsat Services Corporation
• Original Assignee: International Telecommunications Satellite Organization
• Inventors: Agrawal, Brij N., Madon, Pierre J.
• Primary Examiner(s): Barefoot, Galen L.
• Application Number:
  US07/044,327
• Time in Patent Office: 1,062 Days
• Field of Search: 244/164, 244/165, 244/169, 244/171, 364/434
• US Class Current: 244/165
• CPC Class Codes: B64G 1/244 Spacecraft control systems
• CPC Class Codes: B64G 1/28 using inertia or gyro effect
• CPC Class Codes: B64G 1/365 using horizon or Earth sensors

Abstract

A system and method for attitude control in a geosynchronous satellite to compensate for roll and yaw pointing errors consequent to orbit inclination variation from the nominal equatorial orbit plane provides for an inertially fixed momentum vector coupled to the satellite through a gimbal system providing a one and preferably two degree-of-freedom relationship with the momentum vector. In a two degree-of-freedom embodiment, the momentum vector is established by spinning the satellite about an axis or providing an independent momentum wheel with the gimbal axes provided along the roll and yaw axes. Gimbal torquers torque the satellite about the inertially fixed momentum vector in a time-varying manner to effect correction of the roll and yaw pointing errors. Roll and yaw pointing errors consequent to orbit inclination drift from the nominal equatorial orbit are corrected in a fuel-efficient manner to extend the operating life of the satellite.

Claims

1. A method of correcting pointing errors in a geosynchronous satellite consequent to orbit inclination, comprising the steps of:

establishing the momentum vector along an axis other than an axis normal to the equatorial orbit plane;

coupling the momentum vector to the satellite through at least a single degree-of-freedom connection coincident with at least a selected one of the roll or yaw axes for effecting rotation of the satellite relative to the momentum vector about the selected axis, the roll axis tangential to the orbit path and the yaw axis perpendicular to the roll axis in the orbit plane; and

torquing the satellite about the selected axis in a time-varying manner in synchronism with the orbit of the satellite and as a function of an orbit inclination angle i.

2. The method of claim 1, wherein said coupling step further comprises:

coupling the momentum vector to the satellite through at least a single degree-of-freedom connection coincident with the roll axis for effecting rotation of the satellite relative to the momentum vector about the roll axis.

3. The method of claim 2, wherein said torquing step further comprises:

torquing the satellite about the roll axis as a function of -θ

SIN nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

4. The method of claim 1, wherein said coupling step further comprises:

coupling the momentum vector to the satellite through at least a single degree-of-freedom connection coincident with the yaw axis for effecting rotation of the satellite relative to the momentum vector about the yaw axis.

5. The method of claim 4, wherein said torquing step further comprises:

torquing the satellite about said yaw axis as a function of i COS nt where t=time with t=0 at the ascending node and n=the orbit angular rate.

6. The method of claim 5, wherein said torquing step further comprises:

establishing the momentum vector along an axis tilted by an angle θ

from the inclined orbit normal; and

torquing the satellite about said yaw axis by (i+θ

) COS nt where t=time with t=0 at the ascending node and n=the orbit angular rate, θ

has a maximum value of 0.178i.

7. The method of claim 1, wherein said coupling step further comprises:

coupling the momentum vector to the satellite through a two degree-of-freedom connection coincident, respectively, with the roll and yaw axes for effecting rotation of the satellite relative to the momentum vector about the roll and yaw axes.

8. The method of claim 7, wherein said torquing step further comprises:

torquing the satellite about the roll axis as a function of -θ

SIN nt and about said yaw axis as a function of i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

9. The method of claim 8, wherein said torquing step further comprises:

establishing the momentum vector along a normal to the inclined orbit plane; and

torquing the satellite about the roll axis by -θ

SIN nt and about said yaw axis by i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

10. A method of correcting pointing errors in a geosynchronous satellite consequent to orbit inclination, the satellite of the type having a spun section that establishes a momentum vector along an axis and a despun section, comprising the steps of:

establishing a momentum vector along an axis other than an axis normal to the equatorial orbit plane;

coupling the momentum vector of the spun section to the despun section through at least a single degree-of-freedom connection coincident with at least a selected one of the roll or yaw axes for effecting rotation of the despun section relative to the momentum vector about the selected axis, the roll axis tangential to the orbit path and the yaw axis perpendicular to the roll axis in the orbit plane; and

torquing the despun section about the selected axis in a time-varying manner in synchronism with the orbit of the satellite and as a function of an orbit inclination angle i.

11. The method of claim 10, wherein said coupling step further comprises:

coupling the momentum vector of the spun section to the despun section through at least a single degree-of-freedom connection coincident with the roll axis for effecting rotation of the despun section relative to the momentum vector about the roll axis.

12. The method of claim 11, wherein said torquing step further comprises:

torquing the despun section about the roll axis as a function of -θ

SIN nt where t=time with t=0 at the ascending node, n=the orbit angular rate and θ

has a maximum value of 0.178i.

13. The method of claim 10, wherein said coupling step further comprises:

coupling the momentum vector of the spun section to the despun section through at least a single degree-of-freedom connection coincident with the yaw axis for effecting rotation of the despun section relative to the momentum vector about the yaw axis.

14. The method of claim 13, wherein said torquing step further comprises:

torquing the despun section about the yaw axis as a function of i COS nt where t=time with t=0 at the ascending node and n=the orbit angular rate.

15. The method of claim 14, wherein said torquing step further comprises:

establishing the momentum vector along an axis tilted by an angle θ

from the inclined orbit normal; and

torquing the satellite about said yaw axis by (i+θ

) COS nt where t=time with t=0 at the ascending node and n=the orbit angular rate, and θ

has a maximum value of 0.178i.

16. The method of claim 10, wherein said coupling step further comprises:

coupling the mometum vector of the spun section to the despun section through a two degree-of-freedom connection coincident, respectively, with the roll and yaw axes for effecting rotation of the despun section relative to the momentum vector about the roll and yaw axes.

17. The method of claim 16, wherein said torquing step further comprises:

torquing the satellite about the roll axis as a function of -θ

SIN nt and about said yaw axis as a function of i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

18. The method of claim 17, wherein said torquing step further comprises:

establishing the momentum vector along a normal to the inclined orbit plane; and

torquing the satellite about the roll axis by -θ

SIN nt and about said yaw axis by i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

19. A method of correcting pointing errors in a geosynchronous satellite consequent to orbit inclination, comprising the steps of:

rotating at least one mass about an axis to establish a momentum vector along the axis, the axis other than an axis normal to the equatorial orbit plane;

coupling the momentum vector to the satellite through at least a single degree-of-freedom connection coincident with at least a selected one of the roll or yaw axes for effecting rotation of the satellite relative to the momentum vector about the selected axis, the roll axis tangential to the orbit path and the yaw axis perpendicular to the roll axis in the orbit plane; and

torquing the satellite about the selected axis in a time-varying manner in synchronism with the orbit of the satellite and as a function of an orbit inclination angle i.

20. The method of claim 19, wherein said coupling step further comprises:

coupling the momentum vector to the satellite through at least a single degree-of-freedom connection coincident with the roll axis for effecting rotation of the satellite relative to the momentum vector about the roll axis.

21. The method of claim 20, wherein said torquing step further comprises:

torquing the satellite about the roll axis as a function of -θ

SIN nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

22. The method of claim 19, wherein said coupling step further comprises:

coupling the momentum vector to the satellite through at least a single degree-of-freedom connection coincident with the yaw axis for effecting rotation of the satellite relative to the momentum vector about the yaw axis.

23. The method of claim 22, wherein said torquing step further comprises:

torquing the satellite about the yaw axis as a function of i COS nt where t=time with t=0 at the ascending node and n=the orbit angular rate.

24. The method of claim 23, wherein said torquing step further comprises:

establishing the momentum vector along an axis tilted by an angle θ

from the inclined orbit normal; and

torquing the satellite about said yaw axis y (i+θ

) COS nt where t=time with t=0 at the ascending node and n=the orbit angular rate, θ

has a maximum value of 0.178i.

25. The method of claim 19, wherein said coupling step further comprises:

coupling the momentum vector to the satellite through a two degree-of-freedom connection coincident, respectively, with the roll and yaw axes for effecting rotation of the satellite relative to the momentum vector about the roll and yaw axes.

26. The method of claim 25, wherein said torquing step further comprises:

torquing the satellite about the roll axis as a function of θ

SIN nt and about said yaw axis as a function of i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

27. The method of claim 26, wherein said torquing step further comprises:

establishing the momentum vector along a normal to the inclined orbit plane; and

torquing the satellite about the roll axis by -θ

SIN nt and about said yaw axis by i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

28. The method of claim 19, wherein the geosynchronous satellite is of the type having an earth sensor control loop for determining the attitude of the satellite relative to the earth and for correcting the attitude in response to the output of the earth sensor, further comprising the step of:

introducing a bias value into the earth sensor control loop to offset the effect of the torquing step.

29. An attitude control system for a geosynchronous satellite, comprising:

means for establishing a momentum vector by rotation of a mass about a spin axis, the axis other than an axis normal to the equatorial orbit plane;

means for coupling the momentum vector to at least a portion of the geosynchronous satellite for controlled relative rotation of the satellite about at least a selected one of the roll axis or yaw axis, the roll axis tangential to the orbit path and the yaw axis perpendicular to the roll axis in the orbit plane; and

means for driving said coupling means to effect controlled rotation of said satellite about the selected axis in a time-varying manner in synchronism with the orbit of the satellite and as a function of the orbit inclination angle i.

30. The attitude control system for a geosynchronous satellite of claim 29, wherein said first-mentioned means comprises a momentum wheel mounted for rotation about a spin axis to develop a momentum vector therealong.

31. The attitude control system for a geosynchronous satellite of claim 29, wherein said first-mentioned means comprises a first portion of the satellite mounted for rotation about a spin axis relative to another despun portion of the satellite.

32. The attitude control system for a geosynchronous satellite of claim 29, wherein said second-mentioned means comprises a single degree-of-freedom gimbal means having an axis coincident with the roll axis for effecting rotation of the satellite relative to the momentum vector about the roll axis.

33. The attitude control system for a geosynchronous satellite of claim 32, wherein said gimbal means is controlled by said third-mentioned means as a function of -θ

SIN nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.187i.

34. The attitude control system for a geosynchronous satellite of claim 29, wherein said second-mentioned means comprises a single degree-of-freedom gimbal means having an axis coincident with the yaw axis for effecting rotation of the satellite relative to the momentum vector about the yaw axis.

35. The attitude control system for a geosynchronous satellite of claim 34, wherein said gimbal means is controlled by said third-mentioned means as a function of i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and i=the orbit inclination angle.

36. The attitude control system for a geosynchronous satellite of claim 35, wherein said first-mentioned means establishes the momentum vector along an axis tilted by an angle θ

from the inclined orbit plane normal and said gimbal means is controlled by said third-mentioned means by (i+θ

) COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and i=the orbit inclination angle.

37. The attitude control system for a geosynchronous satellite of claim 29, wherein said second-mentioned means comprises a two degree-of-freedom gimbal means having axes coincident with the roll and yaw axes of the satellite.

38. The attitude control system for a geosynchronous satellite of claim 37, wherein the roll axis is controlled by said third-mentioned means as a function of -θ

SIN nt where t=time with t=0 at the ascending node, n=the orbit angular rate and θ

has a maximum value of 0.178i.

39. The attitude control system for a geosynchronous satellite of claim 38, wherein said first-mentioned means establishes the momentum vector along a normal to the inclined orbit plane and the roll axis is controlled by said third-mentioned means by -(i+θ

) SIN nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and θ

has a maximum value of 0.178i.

40. The attitude control system for a geosynchronous satellite of claim 37, wherein the yaw axis is controlled by said third-mentioned means as a function of i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and i=the orbit inclination angle.

41. The attitude control system for a geosynchronous satellite of claim 40, wherein said first-mentioned means establishes the momentum vector along a normal to the inclined orbit plane and the yaw axis is controlled by said third-mentioned means by i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, and i=the orbit inclination angle.

42. The attitude control system for a geosynchronous satellite of claim 37, wherein the roll axis is controlled by said third-mentioned means as a function of -θ

SIN nt and the yaw axis is controlled by said third-mentioned means as a function of i COS nt where t=time with t=0 at the ascending node, n=the orbit angular rate, i=the orbit inclination angle, and θ

has a maximum value of ratio of the earth'"'"'s radius to the geosynchronous orbit altitude multiplied by the orbit inclination angle.

43. An attitude control system for a geosynchronous satellite of the type having a spun section that establishes a momentum vector and a despun section, comprising:

means defining a two degree of freedom coupling between the spun and despun sections for controlled relative rotation of the spun and depsun sections of the satellite about orthogonal roll and yaw axes defined by the two degrees of freedom coupling, the roll axis tangential to the orbit path and the yaw axis perpendicular to the roll axis in the orbit plane; and

means for controlling said coupling means to effect controlled relative rotation between said spun and despun sections in a time-varying manner in synchronism with the orbit of the satellite about the earth and as a function of the orbit inclination angle i.

44. The attitude control system of claim 43, wherein the momentum vector is established along an axis other than an axis normal to the equatorial orbit plane.

45. The attitude control system of claim 44, wherein said second-mentioned means controls said coupling means to provide relative rotation about said roll axis as a function of -θ

SIN nt and relative rotation about said yaw axis as a function of i COS nut where t=time with t=0 at the ascending node, n=the orbit angular rate, i=the orbit inclination angle, and θ

having a maximum value of the ratio of the earth'"'"'s radius to the geosynchronous orbit altitude multiplied by the orbit inclination angle.

46. The attitude control system of claim 44, wherein said second-mentioned means controls said coupling means to provide relative rotation about said roll axis as a function of -θ

SIN nt, where t=time with t=0 at the ascending node, n=the orbit angular rate, i=the orbit inclination angle, and θ

has a maximum value of 0.178i.

47. The attitude control system of claim 44 wherein said second-mentioned means controls said coupling means to provide relative rotation about said yaw axis as a function of i COS nt with t=time with t=0 at the ascending node, n=the orbit angular rate, i=the orbit inclination angle.