Spacecraft payload orientation steering
View Patent ↗Spacecraft payload orientation steering is provided for an orbiting spacecraft in motion along an orbit track around a celestial body, the orbit track having a nominal inclination with respect to an equatorial orbit, a substantial eccentricity, and a drift angle with respect to the nominal inclination. Coordinates of an optimal payload target location as a function of a spacecraft position along the orbit track are determined, the target location being on the surface of the celestial body and having a substantial motion with respect to the surface and with respect to a spacecraft nadir. A payload of the spacecraft is substantially aligned with the determined coordinates by steering the satellite body to correct for at least one of the inclination drift angle, and the eccentricity, thereby adjusting the spacecraft orientation as a function of the spacecraft position along the orbit track.
1. A method comprising:
determining, as a function of a position of a spacecraft along an orbit track around a celestial body, time varying coordinates of a non-fixed target location, the non-fixed target location having a substantial motion with respect to a fixed fit-target on a surface of the celestial body and with respect to a spacecraft nadir, said orbit track having one or both of (i) an eccentricity of at least 0.2, and (ii) an inclination with respect to an equatorial orbit of at least 50 degrees; and
substantially aligning an antenna reflector of the spacecraft with the determined time varying coordinates by actively steering the spacecraft orientation as a function of the position of the spacecraft along the orbit track such that a boresight of the antenna reflector points to the determined time varying coordinates, wherein actively steering the spacecraft orientation includes:
determining an actual orbit position of the spacecraft;
computing, for the determined actual orbit position, a desired inertial attitude of the spacecraft in terms of fitted expressions related to payload optimized roll angle, Φ FIT , pitch angle, θ FIT , and yaw angle, Ψ FIT , wherein
one or more of Φ FIT , θ FIT , and Ψ FIT are expressed as a sum of a first function that computes a steering angle to the fixed fit-target and a respective second or third function modeling a fitted profile away from the fit-target.
2. The method of claim 1 , wherein
the first function computes the angle to the fixed fit-target via analytical geometry;
the second function is a multidimensional high order polynomial function modeling a fitted profile away from the fit-target; and
the third function is an m th order Fourier series modeling a fitted profile away from the fit-target.
3. The method of claim 1 , wherein the orbit track has an eccentricity of at least 0.24, and an inclination with respect to the equatorial orbit of at least 50 degrees.
4. The method of claim 3 , wherein an orbital inclination is permitted to drift over time.
5. The method of claim 1 , wherein determining the coordinates is adapted to optimize EIRP.
6. The method of claim 1 , wherein the spacecraft is steered in yaw.
7. The method of claim 1 , wherein determining the coordinates of the target location is autonomously performed by logic embedded in a control electronics module of the spacecraft.
8. The method of claim 7 , wherein any drift in the orbit is automatically predicted by an on-board orbit propagator processor embedded in the control electronics module of the spacecraft.
9. The method of claim 1 , wherein the target location is computed using a multidimensional polynomial of Nth order, parameterized by the spacecraft position, describing at least one of a roll, pitch, and yaw rotation away from a predetermined fixed target location, with N being greater than or equal to 1.
10. The method of claim 1 , wherein the target location is computed using an Mth order Fourier series, parameterized by the spacecraft position, describing at least one of a roll, pitch, and yaw rotation away from a predetermined fixed target location, with M being greater than or equal to 1.
11. The method of claim 1 , wherein Euler angles are used to represent subsequent angular rotations.
12. The method of claim 1 , wherein quaternions are used to represent the required angular rotations and reference frames.
13. A system for spacecraft payload orientation steering, said system comprising:
an orbit propagator processor configured to determine, as a function of a position of a spacecraft along an orbit track around a celestial body, time varying coordinates of a non-fixed target location, the non-fixed target location having a substantial motion with respect to a fixed fit-target on a surface of the celestial body and with respect to a spacecraft nadir, said orbit track having one or both of (i) an eccentricity of at least 0.2 and (ii) an inclination with respect to an equatorial orbit of at least 50 degrees; and
an attitude control subsystem programmed to substantially align an antenna reflector of the spacecraft with the determined time varying coordinates by actively steering the spacecraft orientation as a function of the position of the spacecraft along the orbit track such that a boresight of the antenna reflector points to the determined time varying coordinates, wherein actively steering the spacecraft orientation includes:
determining an actual orbit position of the spacecraft;
computing, for the determined actual orbit position, a desired inertial attitude of the spacecraft in terms of fitted expressions related to payload optimized roll angle, Φ FIT , pitch angle, ⊖ FIT , and yaw angle, Ψ FIT , wherein one or more of Φ FIT , ⊖ FIT and Ψ FIT are expressed as a sum of a first function that computes a steering angle to the fixed fit-target and a respective second or third function modeling a fitted profile away from the fit-target.
14. The system of claim 13 , wherein:
the first function computes the angle to the fixed fit-target via analytical geometry;
the second function is a multidimensional high order polynomial function modeling a fitted profile away from the fit-target; and
the third function is an m th order Fourier series modeling a fitted profile away from the fit-target.
15. The system of claim 13 , wherein the orbit track has an eccentricity of at least 0.24, and an inclination with respect to the equatorial orbit of at least 50 degrees.
16. The system of claim 15 , wherein an orbital inclination is permitted to drift over time.
17. The system of claim 13 , wherein determining the coordinates is adapted to optimize EIRP.
18. The system of claim 13 , wherein the spacecraft is steered in yaw.
19. The system of claim 13 , wherein determining the coordinates of the target location is autonomously performed by logic embedded in a control electronics module of the spacecraft.
20. The system of claim 19 , wherein any drift in the orbit is automatically predicted by an on-board orbit propagator processor embedded in the control electronics module of the spacecraft.
21. The system of claim 13 , wherein the target location is computed using a multidimensional polynomial of Nth order, parameterized by the spacecraft position, describing at least one of a roll, pitch, and yaw rotation away from a predetermined fixed target location, with N being greater than or equal to 1.
22. The system of claim 13 , wherein the target location is computed using an Mth order Fourier series, parameterized by the spacecraft position, describing at least one of a roll, pitch, and yaw rotation away from a predetermined fixed target location, with M being greater than or equal to 1.
23. The system of claim 13 , wherein Euler angles are used to represent subsequent angular rotations.
24. The system of claim 13 , wherein quaternions are used to represent the required angular rotations and reference frames.