Method and system for controlling the orbit of a satellite in earth orbit
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
B64G-001/24
B64G-001/26
B64G-001/28
F02K-009/84
출원번호
US-0318670
(2015-06-19)
등록번호
US-10232959
(2019-03-19)
우선권정보
FR-14 55630 (2014-06-19)
국제출원번호
PCT/EP2015/063879
(2015-06-19)
국제공개번호
WO2015/193499
(2015-12-23)
발명자
/ 주소
Moro, Valerio
Fischer, Jean
출원인 / 주소
AIRBUS DEFENCE AND SPACE SAS
대리인 / 주소
Im IP Law
인용정보
피인용 횟수 :
0인용 특허 :
14
초록▼
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-cont
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.
대표청구항▼
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 referen
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: |EN1+EN2+RN·sin(ΔT)|>Γin 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, andΔT 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: ∥F1⊗F2∥>∧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; anddetermining 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; andwherein 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 that in the satellite on the station in the Earth orbit, 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;wherein said at least one transporter is configured 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;a controller comprising a set of executable program code, the code programmed to control simultaneously an inclination of the orbit of the satellite, a longitude of the orbit of the satellite and an angular momentum stored in an angular momentum storage device by commanding the propulsion system and said at least one transporter according to a maneuver plan comprising at least two orbit control maneuvers;thrust forces of the propulsion system during said at least two orbit control maneuvers have respective thrust directions that are not parallel in an inertial frame of reference; andwherein the thrust forces of said at least two orbit control maneuvers are determined to control simultaneously the inclination and the 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. 15. The satellite as claimed in claim 14, wherein said at least one transporter comprises at least two articulated arms arranged on respective opposite sides of the plane XZ formed by the axes X and Z of the satellite frame of reference and non-symmetrically with respect to the plane XZ. 16. The satellite as claimed in claim 14, wherein said at least one 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 to one another for at least two pairs of adjacent articulations. 17. The satellite as claimed in claim 16, wherein the articulated arm comprises an additional articulation having at least one degree of freedom in rotation about its rotation axis. 18. The satellite as claimed in claim 14, wherein the satellite comprises an additional thruster of a fixed orientation relative to the satellite. 19. The satellite as claimed in claim 14, wherein the propulsion system carried by the transporter is an electrical propulsion system. 20. The satellite as claimed in claim 14, 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. 21. An orbit control system for a satellite as claimed in claim 14, further comprising a processor configured to: determine an inclination control requirement for the orbit of the satellite;determine a longitude control requirement for the orbit of the satellite;determine an angular momentum unload requirement for the angular momentum storage device of the satellite;determine the thrust forces of said at least two orbit control maneuvers of the maneuver plan as a function of the inclination control requirement, of the longitude control requirement and of the angular momentum unload requirement, andcontrol, with the determined thrust forces, simultaneously the inclination and the 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. 22. The satellite orbit control system as claimed in claim 21, wherein the maneuver plan verifies a following condition: |EN1+EN2+RN·sin(ΔT)|>Γin 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, andΔT 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. 23. The satellite orbit control system as claimed in claim 21, wherein the maneuver plan verifies a following condition: ∥F1⊗F2∥>∧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. 24. The satellite orbit control system as claimed in claim 21, wherein the processor for determining the maneuver plan comprises a processor of the satellite and a processor of a ground station. 25. The satellite orbit control system as claimed in claim 24, wherein an intermediate maneuver plan configured to control only the orbit of the satellite is determined by the processor of the ground station and transmitted to the satellite, and the maneuver plan to unload the angular momentum from the angular momentum storage device is determined by the processor of the satellite as a function of the intermediate maneuver plan and of the angular momentum unload requirement.
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이 특허에 인용된 특허 (14)
Montenbruck Oliver (Mnchen DEX) Eckstein Martin (Wrthsee DEX) Werner Wilhelm (Bieringen (Schntal) DEX), Apparatus for orbit control of at least two co-located geostationary satellites.
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