[특허]Method and apparatus for solar tacking momentum maintenance in long-duration deployment of a large reflector

The Boeing Company

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특허 상세정보

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Method and apparatus for solar tacking momentum maintenance in long-duration deployment of a large reflector

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	E03H-005/00
미국특허분류(USC)	244/164; 244/158.R
출원번호	US-0318715 (2002-12-13)
발명자 / 주소	Wang,Hanching Grant Fowell,Richard A. Ih,Che Hang C.
출원인 / 주소	The Boeing Company
대리인 / 주소	Gates &
인용정보	피인용 횟수 : 2 인용 특허 : 8

초록

A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method uses solar tacking and similar techniques to produce differential solar torques that are used to control the momentum and attitude of the spacecraft during the appendage deployment.

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What is claimed is: 1. A method of controlling an attitude of a spacecraft while deploying an appendage from the spacecraft, the spacecraft having variable mass properties from the appendage deployment, and having a plurality of reaction wheels and a plurality of solar wings, comprising the steps of: determining a direction and an amount of momentum to add to the spacecraft via the plurality of reaction wheels ({right arrow over (M) }spin); storing the determined momentum in at least one of the reaction wheels to provide momentum in the determined direction and in the determined amount; and deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft. 2. The method of claim 1, wherein the solar wings have a longitudinal axis, and the attitude of the spacecraft is controlled by routing the solar wings about the longitudinal axis. 3. The method of claim 1, wherein the solar wings have a longitudinal axis, and the attitude of the spacecraft is controlled by rotating the solar wings about an axis perpendicular to the longitudinal axis. 4. The method of claim 1, wherein the determined direction of the momentum added to the spacecraft via the reaction wheels is about a sun line axis, and the sun line axis is a spacecraft spin axis. 5. The method of claim 1, wherein the amount of momentum to store in the at least one momentum wheel ∥{right arrow over (M)} spin∥ is such that throughout the deployment of the appendage, wherein I spiline is the spacecraft moment of inertia about a spin axis, I transverse is the spacecraft moment of inertia about an axis perpendicular to the first axis, and ωs is a spin rate of the spacecraft. 6. The method of claim 1, wherein the step of determining a direction and amount of momentum to add to the spacecraft via the plurality of momentum wheels ({right arrow over (M)}spin) comprises the steps of: determining a pre-deployment spacecraft momentum {right arrow over (M)}b; determining a desired post-deployment spacecraft momentum {right arrow over (M)}a; determining a momentum change during deployment {right arrow over (M)}l; and determining {right arrow over (M)}spin from {right arrow over (M)}={right arrow over (M)}a-{right arrow over (M)} b-{right arrow over (M)}l, wherein a spin momentum H s={right arrow over (M)}spin={right arrow over (M)}쨌{right arrow over (u)} and {right arrow over (u)} is the unit vector of the first axis. 7. The method of claim 1, wherein the step of deploying the appendage while controlling the attitude of the spacecraft using or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft comprises the step of: deploying the appendage while applying differential torque to the spacecraft by orienting the solar wings to maintain {right arrow over (M)}spin. 8. The method of claim 1, wherein the step of deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft comprises the step of: estimating an external torque applied by orienting the solar wings using angular acceleration estimates derived from spacecraft angular rate measurements; and further orienting the solar wing based on the estimated angular acceleration. 9. The method of claim 1, wherein the step of deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft comprises the step of: estimating a spacecraft momentum; and orienting the at least one of the solar wings based on the estimated spacecraft momentum. 10. The method of claim 1, wherein the step of deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft comprises the step of: estimating a spacecraft attitude with an estimator using spacecraft attitude measurements; and controlling the spacecraft attitude using one or more reaction wheels based on the estimated spacecraft attitude. 11. The method of claim 1, wherein the step of determining a direction and an amount of momentum to add to the spacecraft via the plurality of reaction wheels comprises the steps of: predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage; and computing the direction and amount of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque. 12. The method of claim 11, wherein the environmental torque includes an outgassing torque Tog component and a solar torque Tsp component. 13. The method of claim 11, wherein the momentum includes a transverse momentum component Ht and the momentum further includes a spin momentum Hs, and wherein the step of computing a magnitude and direction of a momentum to add to the spacecraft to at least partially oppose the predicted environmental torque comprises the steps of: computing the transverse momentum Hs perpendicular to the environmental torque and the spacecraft spin axis; and computing the spin momentum Ht along the spin axis of the spacecraft to render the spacecraft passively stable during deployment of the appendage. 14. An apparatus of controlling an attitude of a spacecraft while deploying an appendage from the spacecraft, the spacecraft having variable mass properties from the appendage deployment, and having a plurality of reaction wheels and a plurality of solar wings, comprising: means for determining a direction and an amount of momentum ({right arrow over (M)}spin) to add to the spacecraft via the plurality of reaction wheels; means for storing the momentum in at least one of the reaction wheels to provide momentum in the determined direction and in the determined amount; and means for deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft. 15. The apparatus of claim 14, wherein the solar wings have a longitudinal axis, and the attitude of the spacecraft is controlled by rotating the solar wings about the longitudinal axis. 16. The apparatus of claim 14, wherein the solar wings have a longitudinal axis, and the attitude of the spacecraft is controlled by rotating the solar wings about an axis perpendicular to the longitudinal axis. 17. The apparatus of claim 14, wherein the determined direction of the momentum added to the spacecraft via the reaction wheels is about a sun line axis. 18. The apparatus of claim 14, wherein the amount of momentum to store in the at least one reaction wheel ∥{right arrow over (M)} spin∥ is such that throughout the deployment of the appendage, wherein I spinline is the spacecraft moment of inertia about a spin axis, I t is the spacecraft moment of inertia about an axis perpendicular to the first axis, and ωs is a spin rate of the spacecraft. 19. The apparatus of claim 14, wherein the means for determining a direction and amount of momentum to add to the spacecraft via the plurality of reaction wheels ({right arrow over (M)}spin) comprises: means for determining a pre-deployment spacecraft momentum {right arrow over (M)}b; means for determining a desired post-deployment spacecraft momentum {right arrow over (M)}a; means for determining a momentum change during deployment {right arrow over (M)}l; and means for determining {right arrow over (M)}spin from {right arrow over (M)}={right arrow over (M)}a -{right arrow over (M)}b -{right arrow over (M)}l, wherein a spin momentum Hs ={right arrow over (M)}spin ={right arrow over (M)}쨌{right arrow over (u)} and {right arrow over (u)} is the unit vector of a spin axis. 20. The apparatus of claim 14, wherein the means for deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft comprises: means for deploying the appendage while applying differential torque to the spacecraft by orienting the solar wings to maintain {right arrow over (M)}spin. 21. The apparatus of claim 14, wherein the means for deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft by orienting at least one of the solar wings to apply differential solar torque to the spacecraft comprises: means for estimating an external torque applied by orienting the solar wings using angular acceleration estimates derived from spacecraft angular rate measurements; and means for further orienting the solar wing based on the estimated angular acceleration. 22. The apparatus of claim 14, wherein the means for deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft at least in part by orienting the solar wings to apply differential solar torque to the spacecraft comprises: means for estimating a spacecraft momentum; and means for orienting the at least one of the solar wings based on the estimated spacecraft momentum. 23. The apparatus of claim 14, wherein the means for deploying the appendage while controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft at least in pan by orienting the solar wings to apply differential solar torque to the spacecraft comprises: means for estimating a spacecraft attitude with an estimator using spacecraft attitude measurements; and means for controlling the spacecraft attitude using one or more reaction wheels based on the estimated spacecraft attitude. 24. The apparatus of claim 14, wherein the means for determining a direction and an amount of momentum to add to the spacecraft via the plurality of reaction wheels comprises: means for predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage; and means for computing the direction and amount of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque. 25. The apparatus of claim 24, wherein the environmental torque includes an outgassing torque Tog component and a solar torque Tsp component. 26. The apparatus of claim 24, wherein the momentum includes a transverse momentum component Ht and the momentum further includes a spin momentum Hs, and wherein the means for computing a magnitude and direction of a momentum to add to the spacecraft to at least partially oppose the predicted environmental torque comprises: means for computing the transverse momentum Hs perpendicular to the environmental torque and the spacecraft spin axis; and means for computing the spin momentum Ht along the spin axis of the spacecraft to render the spacecraft passively stable during deployment of the appendage. 27. An apparatus of controlling an attitude of a spacecraft while deploying an appendage from the spacecraft; the spacecraft having substantially variable mass properties from the appendage deployment, and a plurality of solar wings, comprising: a processor for determining a direction and an amount of momentum ({right arrow over (M)}spin) to add to the spacecraft; a plurality of reaction wheels for storing the determined direction and amount of momentum; an appendage driver for deploying the appendage; and a solar wing driver for controlling the attitude of the spacecraft using one or more of the reaction wheels and controlling the momentum of the spacecraft while deploying the appendage by orienting at least one of the solar wings to apply differential solar torque to the spacecraft. 28. The apparatus of claim 27, wherein the solar wings have a longitudinal axis, and the attitude of the spacecraft is controlled by rotating the solar wings about the longitudinal axis. 29. The apparatus of claim 27, wherein the solar wings have a longitudinal axis, and the attitude of the spacecraft is controlled by rotating the solar wings about an axis perpendicular to the longitudinal axis. 30. The apparatus of claim 27, wherein the determined direction of the momentum added to the spacecraft via the reaction wheels is about the sun line axis. 31. The apparatus of claim 27, wherein the amount of momentum to store in the at least one reaction wheel ∥{right arrow over (M)} spin∥ is such that throughout the deployment of the appendage, wherein I spinline is the spacecraft moment of inertia about first axis, I t is the spacecraft moment of inertia about an axis perpendicular to the first axis, and ωs is a spin rate of the spacecraft. 32. The apparatus of claim 27, wherein: the processor further estimates the external torque applied by orienting the solar wing using angular acceleration estimates derived from spacecraft angular rate measurements; and the solar wing driver further orients the solar wings based on the estimated angular acceleration. 33. The apparatus of claim 27, wherein: the processor further estimates a spacecraft momentum; and the solar wing driver further orients the solar wing based on the estimated spacecraft momentum. 34. The apparatus of claim 27, wherein: the processor further estimates the spacecraft attitude with an estimator using spacecraft attitude measurements; and the spacecraft attitude is controlled at least in part using reaction wheels based on the estimated spacecraft attitude. 35. The method of claim 1, wherein the at least one of the solar wings is oriented to prevent saturation of the at least one of the reaction wheels. 36. The apparatus of claim 14, wherein the at least one of the solar wings is oriented to prevent saturation of the at least one of the reaction wheels. 37. The apparatus of claim 27, wherein the at least one of the solar wings is oriented to prevent saturation of the at least one of the reaction wheels.

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이 특허를 인용한 특허 피인용횟수: 2

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Method and apparatus for solar tacking momentum maintenance in long-duration deployment of a large reflector

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Method and apparatus for solar tacking momentum maintenance in long-duration deployment of a large reflector

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