Method and system for controlling the orbit of a satellite in earth orbit

Patent metadata

Field	Value
Publication number	US-10232959-B2
Application number	US-201515318670-A
Country	US
Kind code	B2
Filing date	Jun 19, 2015
Priority date	Jun 19, 2014
Publication date	Mar 19, 2019
Grant date	Mar 19, 2019

Abstract

A method for controlling the orbit of a satellite in earth orbit. The orbit of the satellite is controlled by commanding, according to a maneuver plan, a propulsion system having at least one thruster and a transporter to move the propulsion system. The maneuver plan includes at least two orbit-control maneuvers. The thrust powers of the propulsion system during the two orbit control maneuvers have respective thrust directions that are not parallel in an inertial frame of reference. Each thrust power is determined to simultaneously control the inclination and the position of the orbit of the satellite as well as to form a momentum that is suitable for unloading a device for storing angular momentum of the satellite in a plane orthogonal to the direction of thrust of the thrust power.

First claim

The invention claimed is:

1. A method for controlling an orbit of a satellite on station in an Earth orbit, the satellite comprising an angular momentum device, a propulsion system comprising at least one thruster and a satellite frame of reference centered on a center of mass of the satellite, the satellite frame of reference includes three axes X, Y and Z, the axis X is parallel to a speed vector of the satellite, the axis Z is directed toward the Earth, and the axis Y is orthogonal to the axes X and Z, the method comprises steps of:
determining an inclination control requirement for the orbit of the satellite;
determining a longitude control requirement for the orbit of the satellite;
determining an angular momentum unload requirement for the angular momentum storage device of the satellite;
determining a maneuver plan comprising at least two orbit control maneuvers with respective thrust forces of the propulsion system having respective thrust directions that are not parallel in an inertial frame of reference;
determining the thrust forces as a function of the inclination control requirement, of the longitude control requirement and of the angular momentum unload requirement, wherein the thrust forces of said at least two orbit control maneuvers are determined to control simultaneously an inclination and a longitude of the orbit of the satellite while producing torques configured to unload the angular momentum storage device of the satellite in respective planes that are not parallel in the inertial frame of reference, so that said at least two orbit control maneuvers of the maneuver plan unloads the angular momentum storage device about the three axes;
controlling simultaneously the inclination and the longitude of the orbit of the satellite and an angular momentum stored in the angular momentum storage device by commanding the propulsion system and a transporter of the propulsion system to apply the thrust forces of said at least two orbit control maneuvers of the maneuver plan;
wherein the transporter is commanded to:
modify angles between a thrust direction of said at least one thruster and the axes X, Y of the satellite frame of reference, respectively;
move said at least one thruster while maintaining a constant thrust direction in the satellite frame of reference to produce a torque about any axis in a plane orthogonal to said thrust direction.

2. The method as claimed in claim 1 , wherein the maneuver plan is determined to provide a predetermined minimum unloading capacity about the three axes throughout said at least two orbit control maneuvers of the maneuver plan.

3. The method as claimed in claim 1 , wherein the maneuver plan verifies a following condition:
| EN 1+ EN 2+ RN ·sin(Δ T )|>Γ

in which expression:
Γ is a strictly positive scalar value,
EN1 corresponds to a ratio between a component along the axis X and a component along the axis Y of a thrust force of a first orbit control maneuver of the maneuver plan,
EN2 corresponds to a ratio between a component along the axis X and a component along the axis Y of a thrust force of a second orbit control maneuver of the maneuver plan,
RN corresponds to a ratio between a component along the axis Z and the component along the axis Y of the thrust force of the first or second orbit control maneuver of the maneuver plan, and
ΔT is equal to 2π·(T2−T1−Torb/2)/Torb, in which expression T1 and T2 are times of the first and second orbit control maneuvers and Torb is an orbital period of the satellite.

4. The method as claimed in claim 1 , wherein the maneuver plan verifies a following condition:
∥ F 1⊗ F 2∥>∧

in which expression:
∧ is a strictly positive scalar value, and
∥F1⊗F2∥ corresponds to a norm of a cross product between the thrust forces F1 and F2 of two orbit control maneuvers of the maneuver plan.

5. The method as claimed in claim 1 , wherein the thrust forces of said at least two orbit control maneuvers are not aligned in the satellite frame of reference and times of said at least two orbit control maneuvers have a temporal spacing different from half of an orbital period of the satellite.

6. The method as claimed in claim 1 , wherein the transporter comprises an articulated arm carrying said at least one thruster of the propulsion system, the articulated arm comprises at least three articulations, each articulation having at least one degree of freedom in rotation about a rotation axis, respective rotation axes of adjacent articulations are not parallel for at least two pairs of adjacent articulations, a thrust force of said at least one thruster is controlled by commanding the articulations of the articulated arm.

7. The method as claimed in claim 6 , wherein the articulated arm comprises at least one additional articulation, during at least one orbit control maneuver of the maneuver plan, wherein the method further comprises steps of:
determining an eccentricity control requirement for the orbit of the satellite; and
determining the thrust forces of said at least two orbit control maneuvers of the maneuver plan as a function of the eccentricity control requirement to additionally control the eccentricity of the orbit of the satellite during the maneuver plan.

8. The method as claimed in claim 6 , wherein an eccentricity of the orbit of the satellite is controlled by commanding an additional thruster of the satellite of a fixed orientation relative to the satellite.

9. The method as claimed in claim 6 , wherein at least one of times and durations of said at least two orbit control maneuvers of the maneuver plan are determined to control an eccentricity of the orbit of the satellite during the maneuver plan.

10. The method as claimed in claim 1 , wherein the transporter comprises an attitude control device of the satellite and an articulated arm carrying said at least one thruster of the propulsion system, the articulated arm comprises at least two articulations, each articulation having at least one degree of freedom in rotation, wherein commanding the transporter to apply the thrust forces of said at least two orbit control maneuvers of the maneuver plan comprises: commanding the articulations of the articulated arm and modifying the attitude of the satellite by the attitude control device.

11. The method as claimed in claim 1 , further comprising:
determining, by a ground station and as a function of the inclination orbit control requirement and of the longitude orbit control requirement, an intermediate maneuver plan configured to control the inclination and the longitude of the orbit of the satellite without controlling the angular momentum stored in the angular momentum storage device;
transmitting, by the ground station, the intermediate maneuver plan to the satellite; and
wherein the maneuver plan is determined by the satellite as a function of the intermediate maneuver plan received and of the angular momentum unload requirement.

12. The method as claimed in claim 1 , wherein the maneuver plan comprises at most two orbit control maneuvers per orbital period of the satellite.

13. A non-transitory computer readable medium comprising a set of executable program code, the code programs a processor to be configured to execute the method as claimed in claim 1 for controlling the orbit of the satellite.

14. A satellite to be placed at a station in an Earth orbit, comprising:
a propulsion system comprising at least one thruster and at least one transporter to move the propulsion system in a satellite frame of reference centered on a center of mass of the satellite, the satellite frame of reference includes three axes X, Y and Z such th

Assignees

• Airbus Defence & Space Sas

Inventors

• Moro Valerio
• Fischer Jean

Classifications

• B64G1/283 Primary

  using reaction wheels · CPC title

• F02K9/84

  using movable nozzles · CPC title

• B64G1/286

  using control momentum gyroscopes (CMGs) · CPC title

• B64G1/26

  using jets · CPC title

• B64G1/242 Primary

  Orbits and trajectories · CPC title