Patent Document ID: 9776741
Application ID: 15371579
Patent Flag: 1

Claim One:
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 Σ 1 , converting the flexible spacecraft dynamical system Σ 1 into a flexible spacecraft dynamical system Σ 2 , and incorporating spacecraft rigid-flexible coupling dynamic disturbance into the flexible spacecraft dynamical system Σ 2 ; b) constructing an external system Σ 3 , and describing the rigid-flexible coupling dynamic disturbance through the external system Σ 3 ; the rigid-flexible coupling dynamic disturbance d 0 is expressed as d 0 =F(C d {dot over (η)}+Λη), in which F is a rigid-flexible coupling matrix of a flexible appendage and a body, C d 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 Σ 3 describing the rigid-flexible coupling dynamic disturbance d 0 as: Σ 3 : { w. = Ww + H ⁡ ( u + d 1 ) d 0 = Vw ⁢ ⁢ in ⁢ ⁢ which , w ⁡ [ η T η. T ] T , W = [ 0 I - M - 1 ⁢ Λ - M - 1 ⁢ C d ] , ⁢ H = [ 0 - 1 J ⁢ M - 1 ⁢ F T ] , V = [ F ⁢ ⁢ Λ FC d ] w is a disturbance state variable of the external system Σ 3 , {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 = [ θ θ. ] T ; (2) converting the system Σ 2 into a state space system Σ 4 , which is expressed as: Σ 4 : { x. = Ax + B ⁡ ( u + d 0 + d 1 ) y = Cx ⁢ ⁢ wherein , A = [ 0 1 0 0 ] , B = [ 0 ( J - FF T ) - 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: { d 0 ^ = V ⁢ w ^ w ^ = v - Ly v. = ( W + LVBV ) ⁢ w ^ + ( H + LCB ) ⁢ u wherein, {circumflex over (d)} 0 is an estimated value of the rigid-flexible coupling dynamic disturbance d 0 , 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 Σ 5 of the disturbance controller is expressed as: 
 Σ 5 :ė w =( W+LCBV ) e w +LVAx +( LVB+H ) d 1 wherein, e w =w−ŵ, ŵ is an estimated value of the disturbance state variable w and ė w is a derivative of e w ; d) configuring a dynamic output feedback H ∞ controller; wherein the dynamic output feedback H ∞ controller in the step d) is expressed as: { x. k = A k ⁢ x k + B k ⁢ y u 1 = C k ⁢ x k + D k ⁢ y wherein, u 1 is input of the dynamic output feedback H ∞ controller, x k is a controller state, A x , B x , C x and D x 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 Σ 6 , wherein Σ 6 is expressed as: Σ 6 : { x. k = A k ⁢ x k + B k ⁢ y u = C k ⁢ x k + D k ⁢ y - d 0 ^ wherein u is the control input, and {circumflex over (d)} 0 is the value of the rigid-flexible coupling dynamic disturbance d 0 estimated by the disturbance observer, and the flexible spacecraft refined attitude control system Σ 6 is used to compensate for the rigid-flexible coupling dynamic disturbance through the estimated value.