[특허]Gyroless three-axis sun acquisition using sun sensor and unscented kalman filter

The Boeing Company

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특허 상세정보

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내보내기 페이스북 공유 트위터 공유 카카오톡 공유

Gyroless three-axis sun acquisition using sun sensor and unscented kalman filter

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	B64G-001/36
출원번호	US-0181492 (2014-02-14)
등록번호	US-9643740 (2017-05-09)
발명자 / 주소	Tsao, Tung-Ching Chiang, Richard Y.
출원인 / 주소	The Boeing Company
대리인 / 주소	Ostrager Chong Flaherty & Broitman P.C.
인용정보	피인용 횟수 : 0 인용 특허 : 9

초록

An attitude estimator that uses sun sensor outputs as the only attitude determination measurements to provide three-axis attitude information. This is accomplished by incorporating the Euler equation into the estimator. An unscented Kalman filter is employed to accommodate various nonlinear characteristics and uncertainties of the spacecraft dynamics and thus improve the robustness and accuracy of the attitude estimate.

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1. A method of estimating a three-axis attitude of a spacecraft, comprising: (a) measuring sun unit vectors at successive instants of time using a plurality of sun sensors attached to a spacecraft;(b) calculating angular rates and angular accelerations of the spacecraft about first and second axes perpendicular to the sun line based on the sun unit vector measurements;(c) calculating an angular acceleration of the spacecraft about the sun line using the sun unit vector measurements, a Euler equation and an unscented Kalman filter; and(d) calculating a three-axis attitude of a spacecraft based on the calculation results in steps (b) and (c). 2. The method as recited in claim 1, wherein an estimator state of the unscented Kalman filter consists of a sun line spacecraft angular rate and a sun pointing inertial frame to spacecraft body frame quaternion. 3. The method as recited in claim 1, further comprising using a spacecraft total angular momentum vector as an attitude reference. 4. The method as recited in claim 1, further comprising calculating a spacecraft total angular momentum vector in a sun pointing inertial frame each time an estimate state is updated and adjusting an estimated sun pointing inertial frame to spacecraft body frame quaternion to maintain a fixed orientation of the total angular momentum. 5. The method as recited in claim 3, wherein a reaction wheel offset angular momentum is selected such that when a major or minor axis of the spacecraft is aligned with the spacecraft angular momentum vector, solar panels coupled to the spacecraft will receive maximum current periodically at an angular rate. 6. A spacecraft comprising a body, at least one solar panel coupled to said body, a plurality of sun sensors arranged on said body, and a computer system programmed with software configured to perform the following operations: (a) measuring sun unit vectors at successive instants of time using a plurality of sun sensors attached to a spacecraft;(b) calculating angular rates and angular accelerations of the spacecraft about first and second axes perpendicular to the sun line based on the sun unit vector measurements;(c) calculating an angular acceleration of the spacecraft about the sun line using the sun unit vector measurements, a Euler equation and an unscented Kalman filter; and(d) calculating a three-axis attitude of a spacecraft based on the calculation results in operations (b) and (c). 7. The system as recited in claim 6, wherein an estimator state of said unscented Kalman filter consists of a sun line spacecraft angular rate and a sun pointing inertial frame to spacecraft body frame quaternion. 8. The system as recited in claim 6, further comprising using a spacecraft total angular momentum vector as an attitude reference. 9. The system as recited in claim 6, wherein said computer system is further programmed to calculate a spacecraft total angular momentum vector in a sun pointing inertial frame each time an estimate state is updated and adjusting an estimated sun pointing inertial frame to spacecraft body frame quaternion to maintain a fixed orientation of the total angular momentum. 10. The system as recited in claim 9, wherein a reaction wheel offset angular momentum is selected such that when a major or minor axis of the spacecraft is aligned with the spacecraft angular momentum vector, solar panels coupled to the spacecraft will receive maximum current periodically at an angular rate. 11. A method of commanding a satellite to adjust its attitude to achieve sun acquisition, the method comprising: (a) measuring sun unit vectors at successive instants of time using a plurality of sun sensors attached to a spacecraft;(b) calculating angular rates and angular accelerations of the spacecraft about first and second axes perpendicular to the sun line based on the sun unit vector measurements;(c) calculating an angular acceleration of the spacecraft about the sun line using the sun unit vector measurements, a Euler equation and an unscented Kalman filter;(d) calculating a three-axis attitude of a spacecraft based on the calculation results in steps (b) and (c); and(e) commanding changes in spacecraft attitude based on differences between a target three-axis attitude and said calculated three-axis rates. 12. The method as recited in claim 11, wherein an estimator state of the unscented Kalman filter consists of a sun line spacecraft angular rate and a sun pointing inertial frame to spacecraft body frame quaternion. 13. The method as recited in claim 11, further comprising using a spacecraft total angular momentum vector as an attitude reference. 14. The method as recited in claim 11, further comprising calculating a spacecraft total angular momentum vector in a sun pointing inertial frame each time an estimate state is updated and adjusting an estimated sun pointing inertial frame to spacecraft body frame quaternion to maintain a fixed orientation of the total angular momentum. 15. The method as recited in claim 11, wherein a reaction wheel offset angular momentum is selected such that when a major or minor axis of the spacecraft is aligned with the spacecraft angular momentum vector, solar panels coupled to the spacecraft will receive maximum current periodically at an angular rate. 16. A system for controlling the attitude of a spacecraft based on sun sensor measurements and not gyroscope measurements, comprising: a spacecraft body;at least one solar panel coupled to said body;a plurality of sun sensors arranged on said body;a control processor that outputs torque commands; andmeans for producing torques that change the attitude of the spacecraft in response to torque commands from said attitude control processor,wherein said control processor is programmed to perform operations in accordance with software comprising:an attitude estimator for calculating a three-axis attitude and rate of said spacecraft body in a spacecraft body frame of reference based on sun unit vector measurements made by said plurality of sun sensors and a Euler equation; andan attitude controller that outputs a torque command based on a difference between a target three-axis attitude and said calculated three-axis attitude and rate, andwherein said attitude estimator comprises an unscented Kalman filter having an estimator state consisting of a sun line spacecraft angular rate and a sun pointing inertial frame to spacecraft body frame quaternion. 17. The system as recited in claim 16, wherein said torque-producing means comprise reaction wheels. 18. The system as recited in claim 17, wherein said control processor is further programmed to calculate a spacecraft total angular momentum vector in a sun pointing inertial frame each time an estimate state is updated and adjusting an estimated sun pointing inertial frame to spacecraft body frame quaternion to maintain a fixed orientation of the total angular momentum. 19. The system as recited in claim 18, wherein a reaction wheel offset angular momentum is selected such that when a major or minor axis of the spacecraft is aligned with the spacecraft angular momentum vector, solar panels coupled to the spacecraft will receive maximum current periodically at an angular rate.

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Gyroless three-axis sun acquisition using sun sensor and unscented kalman filter

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특허 상세정보

Gyroless three-axis sun acquisition using sun sensor and unscented kalman filter

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이 특허에 인용된 특허 (9)

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