Method for refined attitude control based on output feedback for flexible spacecraft
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
B64G-001/24
G05D-001/08
G06F-017/16
출원번호
US-0371579
(2016-12-07)
등록번호
US-9776741
(2017-10-03)
우선권정보
CN-2016 1 0389633 (2016-06-02)
발명자
/ 주소
Guo, Lei
Qiao, Jianzhong
Zhang, Ran
Zhang, Peixi
Zhang, Dafa
출원인 / 주소
BEIHANG UNIVERSITY
대리인 / 주소
Locke Lord LLP
인용정보
피인용 횟수 :
0인용 특허 :
2
초록▼
The present invention provides a method for refined attitude control based on output feedback for a flexible spacecraft. The control method comprises the following steps of: a) building a flexible spacecraft dynamical system Σ1, converting the flexible spacecraft dynamical system Σ1 into a flexible
The present invention provides a method for refined attitude control based on output feedback for a flexible spacecraft. The control method comprises the following steps of: a) building a flexible spacecraft dynamical system Σ1, converting the flexible spacecraft dynamical system Σ1 into a flexible spacecraft dynamical system Σ2, and incorporating spacecraft rigid-flexible coupling dynamic disturbance into the flexible spacecraft dynamical system Σ2; b) constructing an external system Σ3, and describing the rigid-flexible coupling dynamic disturbance through the external system Σ3; c) configuring a disturbance observer for estimating the value of the rigid-flexible coupling dynamic disturbance; d) configuring a dynamic output feedback H∞ controller; e) compounding the disturbance observer in step c) with the dynamic output feedback H∞ controller in step d) to obtain a flexible spacecraft refined attitude control system Σ6; the flexible spacecraft refined attitude control system Σ6 compensating for the rigid-flexible coupling dynamic disturbance through the estimated value.
대표청구항▼
1. A method for refined attitude control based on output feedback for a flexible spacecraft, comprising the following steps of: a) building a flexible spacecraft dynamical system Σ1, converting the flexible spacecraft dynamical system Σ1 into a flexible spacecraft dynamical system Σ2, and incorporat
1. A method for refined attitude control based on output feedback for a flexible spacecraft, comprising the following steps of: a) building a flexible spacecraft dynamical system Σ1, converting the flexible spacecraft dynamical system Σ1 into a flexible spacecraft dynamical system Σ2, and incorporating spacecraft rigid-flexible coupling dynamic disturbance into the flexible spacecraft dynamical system Σ2;b) constructing an external system Σ3, and describing the rigid-flexible coupling dynamic disturbance through the external system Σ3; the rigid-flexible coupling dynamic disturbance d0 is expressed as d0=F(Cd{dot over (η)}+Λη), in which F is a rigid-flexible coupling matrix of a flexible appendage and a body, Cd is a modal damping matrix, Λ is a rigidity matrix, η is a mode of the flexible appendage, and {dot over (η)} is a derivative of the mode {dot over (η)} of the flexible appendage;the external system Σ3 describing the rigid-flexible coupling dynamic disturbance d0 as: Σ3:{w.=Ww+H(u+d1)d0=Vwinwhich,w[ηTη.T]T,W=[0I-M-1Λ-M-1Cd],H=[0-1JM-1FT],V=[FΛFCd]w is a disturbance state variable of the external system Σ3, {dot over (w)} is a derivative of w, W, H and V are defined coefficient matrixes, and I is a unit matrix;c) configuring a disturbance observer for estimating a value of the rigid-flexible coupling dynamic disturbance, comprising; (1) constructing a spacecraft attitude angle input matrix x=[θθ.]T;(2) converting the system Σ2 into a state space system Σ4, which is expressed as: Σ4:{x.=Ax+B(u+d0+d1)y=Cxwherein,A=[0100],B=[0(J-FFT)-1],A and B are coefficient matrixes, y is a measurement output, and C is a measurement matrix;(3) configuring the disturbance observer with the aid of the measurement output y, the disturbance observer being expressed as: {d0^=Vw^w^=v-Lyv.=(W+LVBV)w^+(H+LCB)uwherein, {circumflex over (d)}0 is an estimated value of the rigid-flexible coupling dynamic disturbance d0, v is an auxiliary variable, {circumflex over (v)} is a derivative of the auxiliary variable v, y is the measurement output, and L is a disturbance observer gain matrix;observation error dynamic Σ5 of the disturbance controller is expressed as: Σ5:ėw=(W+LCBV)ew+LVAx+(LVB+H)d1 wherein, ew=w−ŵ, ŵ is an estimated value of the disturbance state variable w and ėw is a derivative of ew;d) configuring a dynamic output feedback H∞ controller; wherein the dynamic output feedback H∞ controller in the step d) is expressed as: {x.k=Akxk+Bkyu1=Ckxk+Dkywherein, u1 is input of the dynamic output feedback H∞ controller, xk is a controller state, Ax, Bx, Cx and Dx are controller parameter matrixes to be determined, and the dynamic output feedback H∞ controller suppresses environmental disturbance;e) compounding the disturbance observer in step c) with the dynamic output feedback H∞ controller in step d) to obtain a flexible spacecraft refined attitude control system Σ6, wherein Σ6 is expressed as: Σ6:{x.k=Akxk+Bkyu=Ckxk+Dky-d0^wherein u is the control input, and {circumflex over (d)}0 is the value of the rigid-flexible coupling dynamic disturbance d0 estimated by the disturbance observer, and the flexible spacecraft refined attitude control system Σ6 is used to compensate for the rigid-flexible coupling dynamic disturbance through the estimated value. 2. The method according to claim 1, wherein the flexible spacecraft dynamical system Σ1 is expressed as: Σ1:{Jθ¨+Fη¨=u+d1η¨+Cdη.+Λη+FTθ¨=0wherein, θ is a spacecraft attitude angle, {umlaut over (θ)} is a second-order derivative of the spacecraft attitude angle θ, J is a spacecraft rotational inertia, F is the rigid-flexible coupling matrix of the flexible appendage and the body, FT is a transposed matrix of the rigid-flexible coupling matrix, u is control input, d1 is environmental disturbance torque, η is the mode of the flexible appendage, {dot over (η)} is the derivative of the mode η of the flexible appendage, {dot over (η)} is a second-order derivative of the mode η of the flexible appendage, Cd is the modal damping matrix, and Λ is the rigidity matrix. 3. The method according to claim 2, wherein the modal damping matrix Cd is expressed as Cd=diag{2ζiωi} (i=1, 2, . . . N), in which N is a number of orders of the mode, ζi is modal damping, ωi is a modal frequency, and the rigidity matrix Λ is expressed as Λ=diag{ωi2} (i=1, 2, . . . N). 4. The method according to claim 1, wherein the system Σ2 is expressed as: Σ2:(J−FFT){umlaut over (θ)}=F(Cd{dot over (η)}+Λη)+u+d1 wherein, {umlaut over (θ)} is the second-order derivative of the spacecraft attitude angle θ, J is the spacecraft rotational inertia, F is the rigid-flexible coupling matrix of the flexible appendage and the body, FT is the transposed matrix of the rigid-flexible coupling matrix, u is the control input, d1 is the environmental disturbance torque, η is the mode of the flexible appendage, {dot over (η)} is the derivative of the mode η of the flexible appendage, Cd is the modal damping matrix, and Λ is the rigidity matrix. 5. The method according to claim 1, wherein in the coefficient matrix W, the matrix M is expressed as M=I−FT J−1F, in which I is the unit matrix. 6. The method according to claim 1, wherein the disturbance controller gain matrix L and the controller parameter matrixes Ax, Bx, Cx, Dx to be determined are solved through a convex optimization algorithm as below: making the systems Σ4, Σ5 and Σ6 simultaneous, and obtaining: [x.xk.ew.]=[A+BDkCBCkBVBkCAk0LCA0W+LCBV][xxkew]+[B0LCB+H]d1solving the following convex optimization problem: [Φ11Φ12Φ13Φ14Φ15Φ16*Φ22Φ23000**-γ2I000***-I00****-I0*****-I]0wherein,Φ11=[P2A+BcCAcSA+BcCAc]+[P2A+BcCAcSA+BcCAc]T,Φ12=[P2BV+(YCA)TSBV],Φ13=[P2BSB],Φ14=[C1T0],Φ15=[P200S],Φ16=[BDcCBCcBDcCBCc],Φ22=Q1W+YCBV+(Q1W+YCBV)T,Φ23=(YCB+Q1H),obtainingunknownparameters:L=Q1-1Y,[AkBkCkDk]=[S-100I][AcBcCcDc]. 7. A spacecraft using the method of claim 1. 8. A spacecraft with the refined attitude control for a flexible spacecraft, comprising a spacecraft shell, an external system module, a disturbance observation module, a dynamic output feedback module, a refined attitude control module, a central processing unit (CPU), a control unit, a spacecraft flexible wing plate and a flexible spacecraft dynamical module, wherein the external system module is used to describe the rigid-flexible coupling dynamic disturbance, and delivers the description result of the rigid-flexible coupling dynamic disturbance to the refined attitude control module;the disturbance observation module is used to estimate the value of the rigid-flexible coupling dynamic disturbance by a disturbance observer;the dynamic output feedback module is used to suppress the environmental disturbance by a dynamic output feedback H∞ controller;the refined attitude control module is combined by the disturbance observation module and the dynamic output feedback module, and is used to compensate for the rigid-flexible coupling dynamic disturbance of the spacecraft with the estimated value of the rigid-flexible coupling dynamic disturbance of the spacecraft;the flexible spacecraft dynamical module is used to incorporate the spacecraft rigid-flexible coupling disturbance into a flexible spacecraft dynamical system;the central processing unit (CPU) reads the data of the refined attitude control module and processes the data;the control unit compensates for the rigid-flexible coupling dynamic disturbance of the spacecraft through the refined attitude control module, and adjusts the attitude of the spacecraft; andthe spacecraft flexible wing plate is unfolded at two ends of the spacecraft shell.
연구과제 타임라인
LOADING...
LOADING...
LOADING...
LOADING...
LOADING...
이 특허에 인용된 특허 (2)
Falangas, Eric T., Method of modeling dynamic characteristics of a flight vehicle.
※ AI-Helper는 부적절한 답변을 할 수 있습니다.