CN107272719A - Hypersonic aircraft attitude motion control method for coordinating based on coordinating factor - Google Patents

Abstract

The invention discloses a hypersonic aircraft attitude motion robust coordination control method based on a coordination factor analysis method, belonging to the field of aircraft attitude control. This method first aims at the strong coupling problem of the attitude motion of the aircraft, and decomposes the coupling of the attitude model into three forms: angular motion coupling, inertial coupling and rudder surface control coupling. Secondly, three coordination factors were designed for the above three coupling forms. Then the attitude system is divided into a fast loop and a slow loop, and robust adaptive controllers for the two loops are designed respectively. Finally, the coordination factor is combined with the robust controller to derive the coordination torque, and the coordination torque is distributed to the rudder surface, and the coordination is realized through the deflection of the rudder surface. The method effectively improves the control efficiency of the rudder surface, especially reduces the deflection times of the rudder surface and saves energy.

Description

Attitude-motion Coordinated Control Method for Hypersonic Vehicle Based on Coordination Factor

technical field

The invention discloses a hypersonic aircraft attitude motion coordination control method based on a coordination factor, and belongs to the technical field of spacecraft attitude control.

Background technique

Hypersonic aircraft has important military strategic significance and is the killer weapon for sky warfare in the 21st century. At present, the research on hypersonic vehicles has received widespread attention from all over the world. However, due to the rapid and long-span flight, it also brings great challenges to the control of the aircraft. Due to the characteristics of strong coupling and strong nonlinearity, hypersonic speed has become a research hotspot at present.

In recent years, many valuable scientific research achievements have been made in the attitude motion control of hypersonic vehicles. Many scholars have carried out corresponding work on the strong coupling problem of aircraft attitude motion. Some scholars proposed an online real-time adaptive attitude controller based on feedback linearization and neural network adaptive technology. However, feedback linearization depends on the accuracy of the system model, so it is difficult to guarantee control accuracy when the model is uncertain. Some scholars have also proposed the method of optimal dynamic inverse control, and applied it to attitude control. However, when there is external disturbance in the system, this method is difficult to ensure good control. Subsequently, many scholars proposed nonlinear control methods. Among them, the sliding mode method is used for aircraft control, and two sliding mode controllers based on single loop and inner loop are designed. The research shows that the sliding mode method has good robustness to parameter uncertainties and external disturbances, but it is difficult to deal with strong Coupling problem. Subsequently, some scholars used the hierarchical control idea to deal with the strong coupling problem of hypersonic attitude motion based on the nonlinear method, and some scholars used the decoupling method to deal with the strong coupling problem, including using the singular value perturbation theory to design the inner and outer loops Decouple the controller. Although the above methods have achieved the coordination of attitude motion to a certain extent and laid the foundation for the follow-up work, due to the complexity of the nonlinear dynamics of hypersonic vehicles, these achievements are not clear for the coordination mechanism to be actively linked, suppressed, or Take advantage of coupling effects. The main problem of strong coupling of hypersonic aircraft attitude motion is reflected in the mutual influence between variables. Some coupling effects are beneficial to the control of the aircraft, and some coupling accumulations are fatal to the control of the aircraft. Long-term coupling accumulation will lead to aircraft Instability of posture. Based on the above analysis, it is necessary to study a new control method to deal with the strong coupling problem of hypersonic vehicle attitude motion.

Contents of the invention

In order to overcome the deficiencies of the existing coordinated control methods, the present invention proposes a hypersonic vehicle attitude-motion coordinated control method based on coordination factors. First, based on the mathematical model of the attitude system, the coupling between attitude motions is decomposed into angular motion coupling, inertial coupling and rudder surface manipulation coupling. Then the coordination factors are designed for the above coupling forms, and the coordination factors are combined with the designed robust controller to derive the coordination torque, and the coordination is realized by distributing the torque into the deflection command of the rudder surface. Finally, the effectiveness of the method is verified by simulation, which proves that this method has a good application prospect.

The present invention adopts following technical scheme for solving its technical problem:

A hypersonic vehicle attitude-motion coordinated control method based on a coordination factor, comprising the following steps:

Step 1) Carry out coupling analysis for the attitude model of the hypersonic vehicle, and decompose the strong coupling problem between attitude motions into three forms: angular motion coupling, inertial coupling, and rudder surface control coupling; based on the three forms of coupling, the corresponding state variables are fed back to the corresponding rudder surface circuit design coordination factor;

Step 2) Based on the time-scale separation principle, decompose the attitude system into a slow loop and a fast loop; design a slow loop robust controller and a fast loop robust controller based on the sliding mode method and projection mapping method respectively;

Step 3) Combining the coordination factor with the robust controller to derive the coordination torque, using the rudder distribution matrix, the coordination torque is distributed into the rudder command, and the coordination deflection of the rudder is used to realize the coordination of the attitude movement.

The specific process of the step 1) is as follows:

Step 1-1), establishing a hypersonic vehicle attitude system mathematical model;

in Represents the derivative of Ω, Ω represents the state variable of the slow loop of the system, Ω=[α,β,μ] ^T , α,β,μ are angle of attack, angle of sideslip, and roll angle respectively; represents the derivative of ω, ω represents the state variable of the fast loop of the system, ω=[p,q,r] ^T , p,q,r are the roll rate, pitch rate, and yaw rate respectively; M _c represents the fast loop torque , M _c ＝g _fδ u, where g _fδ ∈ R ^3×3 is the fast loop control surface allocation matrix of the attitude system, u=[δ _e ,δ _a ,δ _r ] ^T , where δ _e ,δ _a ,δ _r are respectively are left elevator aileron rudder, right elevator aileron rudder and rudder; f _s =[f _α ,f _β ,f _μ ] ^T , f _f =[f _p ,f _q ,f _r ] ^T ,

Among them, M, V are the mass and speed of the aircraft respectively; q is the dynamic pressure; S, c, b are the reference area, reference length, and reference width respectively; C _{L, α} are the lift coefficients caused by the angle of attack α, C _{Y, β} is the lateral coefficient caused by sideslip angle β, g is the gravitational coefficient of the earth, C _{l, β} , C _{l, p} , C _lr is the lift coefficient caused by β, p, r, C _{m, α} , C _m,q are the basic pitching moment coefficient and the pitching moment coefficient caused by q, C _D,α is the drag coefficient caused by the angle of attack α, C _n,β ,C _n,p ,C _n,r r are caused by β, p, the yaw moment coefficient caused by r, X _cg is the distance from the center of mass to the reference moment center, Assign matrix coefficients for fast loops. I _xx is the product of inertia around the x-axis, I _yy is the product of inertia around the y-axis, I _zz is the product of inertia around the z-axis, is the derivative of I _xx , is the derivative of I _yy , is the derivative of I _zz , f _s , f _f , g _s , g _f are the attitude system state matrix, l _aero , m _aero , n _aero represent the three-channel torque;

Step 1-2) Carry out coupling analysis to the established attitude system model, decompose the coupling of attitude attitude into attitude angle motion coupling, inertial coupling, rudder surface control coupling,

1) Attitude Angle Coupling Model:

In the attitude system model established in step 1-1), the coupling relation of the attitude angle is described as:

in Respectively expressed as the attitude angle coupling of the slow loop state variables attack angle, sideslip angle and roll angle;

2) Inertial coupling

In the attitude system model established in step 1-1), the inertial coupling is described as:

where f _p , f _q , f _r represent the inertial coupling caused by p, q, r in the fast loop system;

3) Rudder control coupling

In the attitude system model established in step 1-1), the control coupling of the rudder surface is described as:

Where g _l is the steering coupling of the roll channel, g _m is the steering coupling of the pitch channel, and g _n is the coupling of the yaw channel, are the rolling moment coefficients caused by the right elevator, aileron rudder and rudder, respectively, is the pitching moment coefficient caused by the right elevon rudder, The pitching moment coefficient caused by the rudder, is the yaw moment coefficient caused by the right elevon rudder, yaw moment coefficient caused by rudder;

G _f,δ is the distribution matrix of the rudder surface, are the lateral force coefficient caused by the left elevator, the side force coefficient caused by the right elevator aileron and the side force coefficient caused by the rudder, respectively, are the yaw moment coefficient, pitch moment coefficient, roll moment coefficient and drag coefficient caused by the left elevator, is the drag coefficient caused by the right elevator aileron, is the drag coefficient caused by the rudder, is the lift coefficient caused by the left elevator, is the lift coefficient caused by the rudder, is the yaw moment coefficient caused by the right elevator aileron, is the yaw moment coefficient caused by the rudder, Assign matrix parameters for the pitch channel rudder surface, Assign matrix parameters for the yaw channel rudder surface, Assign matrix parameters for the roll channel rudder surface;

Step 1-3) the design of coordination factor;

For 1) attitude angle coupling in steps 1-2), It is set to 0, expressed as psina-rcosα=0, and r=ptanα is obtained, and r=ptanα is fed back to the rudder loop, so the coordination factor is designed as Where k ₁ >0 is a design parameter, is the first component of the rudder coordination factor;

In steps 1-2), 1) the coupling relationship between attack angle and roll angle is described as p-σr<p≤pcosα+rsinα<p+ρr, and the adjustment factor for feeding β to the rudder loop is designed as:

Where |sina|≠Δ _α , k ₂ , k ₃ design parameters greater than zero, is the second component of the rudder coordination factor;

For 2) inertial coupling in step 1-2), feed β,r to the aileron loop, β,q to the rudder loop, and α,p to the elevator loop to increase the damping torque and stability torque; coordination The factorial design is:

where λ _e , λ _a , are the coordination factors of corresponding state variables fed back to the left elevator, right elevon rudder and rudder, k ₄ , k ₅ , k ₆ , k ₇ , k ₈ , k ₉ are design parameters greater than zero;

For 3) rudder control coupling in steps 1-2), define the coupling degree of aileron rudder and rudder control as:

in,

in, is the incremental coefficient of the three-channel moment coefficient change caused by the rudder deflection, is the three-channel moment coefficient change increment coefficient caused by the right elevon rudder deflection, S _cwr , S _cw are expressed as rudder area and vertical tail area; S, L reference area and reference length; y _cwr is the distance from the rudder surface center to the longitudinal axis Distance; χ is the sweep angle; ξ is the correction factor; satisfy S _cwr ≤ S, S _cw ≤ S, y _cwr ≤ L, 0<ξ≤1, 0<cosχ≤1; n is the relative execution efficiency of the aileron; η _k , are the heel-to-tip ratio and stem-to-span ratio of the exposed wing, respectively; f is the ratio of the exposed wing to the root; the coordination strategy of the rudder surface is designed as δ _a = Eδ _r , 0<E≤1, and E is the coupling degree between the rudder and the right elevon rudder.

The specific process of said step 2) is as follows:

Step 2-1) Introduce uncertain parameter vector

By introducing an uncertain parameter vector, the system model in step 1-1) is expressed as

Among them, D stands for external disturbance,

θ ₁ ＝ _CL,α ,θ ₂ ＝C _Y,β ,θ ₃ ＝[C _Y,β C _L,α ] ^T ,

where C _L,α , C _Y,β , C _Y,β , C _L,α , C _n,β , C _n,p , C _n,r , C _m,α , C _m,q , C _D,α , C _l,β , C _l,r , C _l,p are the aerodynamic parameters of the aircraft.

Step 2-2) Projection Mapping Algorithm

Θ _Ω , Θ _ω are denoted as the bounds of uncertain parameter vectors θ _Ω , θ _ω respectively, expressed as:

Θ _Ω ＝{θ _Ω ∈ R ⁴ |θ _{Ωi min ≤θ Ωi ≤θ Ωi max} ,i=1,…,4},

Θ _ω ＝{θ _ω ∈ R ¹⁷ |θ _{ωi min} ≤θ _ωi ≤θ _{ωi max} ,i=1,...,17}

Among them, θ _Ωi , θ _{Ωi min} , θ _{Ωi max} are the components of the slow loop state variable, the minimum value of the component and the maximum value of the component, θ _ωi , θ _{ωi min} , θ _{ωi max} are the components of the fast loop state variable, the minimum value of the component The maximum value and component; let Expressed as an estimate of θ _Ω , Represents the estimation error, θ _ω is similar to θ _Ω , according to the projection mapping algorithm, the adaptive law of uncertain parameters is

in are fast and slow loop adaptive functions respectively, Γ>0 is a diagonal matrix, τ is an adaptive function, and the projection operator defined as

in θ _imax , θ _imin is the estimated value component of the state variable, the maximum value of the state variable component and the minimum value of the state variable component Obviously, for the adaptive function τ,

in is the adaptive estimated value, θ _min and θ _max are the estimated minimum and maximum values, For the estimated error transposition, Γ ^-1 is the inverse of the parameter diagonal matrix, and Γ is the design parameter;

Step 2-3) Attitude Angle Robust Controller Design

Define the tracking error e ₁ =Ω-Ω _c , the sliding mode function is designed as

Among them, K=diag{K ₁ , K ₂ ,K ₃ }, K _i >0, i=1, 2, 3, which represent the design parameters, σ is the sliding mode function, and the first order derivative of the sliding mode function is obtained

obtain The adaptive law of Among them, κ ₁ ＞0, Γ ₁ ∈R ^4×4 , λ _Ω =diag{λ ₁ ,λ ₂ ,λ ₃ ,λ ₄ }＞0 are the controller parameters, g _s , f ₁ , Ψ are the system coefficient matrix , Ψ ^T is the transpose matrix of Ψ, e ₁ is the error, The slow loop follows the derivative of the instruction, is the estimated value of the slow loop state variable, is the estimation error, δ is the sliding mode function;

Step 2-4) Attitude Angle Rate Controller Design

Define the error e ₂ =ω-ω _c , and derive the error

virtual controller As a disturbance, we get in Ξ ^T is the transposition of the system fast loop system matrix; the attitude angle rate loop control design is in is the inverse of the fast loop system state matrix g _f , Ξ ^T is the transposition of the system fast loop system matrix, θ _ω is the fast loop state variable, g _f is the system state matrix, M _c is the control torque, D is the external disturbance, is the derivative of the slow loop controller, κ ₂ is the control design parameter, and is the observed tor output, is the transpose of the slow loop state system matrix, Estimated by the observer, defining the observation error Adaptive law is Among them, κ ₂ ＞0, Γ ₂ ∈R ^17×17 ＞0, λ _ω ∈R ^{17 ×17} ＞0 are the controller parameters, To estimate the error, the observer is designed as

in,

q ₁ (e), q ₂ (e),...,q _n (e) are observer parameters.

The specific process of said step 3) is as follows:

With the attitude angular rate loop controller M _c and the rudder distribution matrix g _fδ , the coordination torque is expressed as M _c =g _Γfδ ·δ, where

in Assign matrix parameters for the pitch channel rudder surface, Assign matrix parameters for the yaw channel rudder surface, Assign matrix parameters to the control surface of the roll channel, λ _e , λ _a , and λ _r are coordination factors; then the control surface command can be calculated as δ=g _Γfδ ^-1 ·M _c .

When the coordination strategy adopts the coordinated deflection of the aileron rudder and rudder to realize the yaw moment and roll moment, and the deflection of the elevator to realize the pitch moment, the coordinated design of the rudder surface is:

δ _coe = δ _e , δ _coa = Eδ _r , δ _cor = δ _r .

The beneficial effects of the present invention are as follows:

1. The performance of the controller with the coordination factor added is better than that of the controller without the coordination factor. The superior performance of the coordinated controller is reflected in the reduced number of vibrations, small overshoot of attitude angle tracking, stable tracking and fast response.

2. The controller adds the coordination factor, and the attitude angle rate becomes stable and fast accordingly.

3. The controller adds a coordination factor, which effectively improves the control efficiency of the rudder surface, especially reduces the deflection times of the rudder surface and saves energy.

Description of drawings

Figure 1 is a schematic diagram of the aircraft attitude movement coordination mechanism.

Figure 2 is the corresponding simulation comparison diagram of the attitude angle with and without coordination factor, where (a) is the comparison diagram of the attack angle tracking curve, (b) is the comparison diagram of the sideslip angle tracking curve, and (c) is the roll angle tracking curve Comparison graph.

Figure 3 is a simulation comparison diagram of attitude angular rate with coordination factor and without coordination factor, in which (a) is a comparison diagram of roll angular rate response, and (b) is a comparison diagram of pitch angular rate response. (c) is a comparison chart of yaw rate response.

Figure 4 is a comparison curve of rudder surface deflection with coordination factor and without coordination factor, (a) is the comparison curve of aileron rudder deflection, (b) is the comparison curve of elevator deflection, and (c) is the comparison curve of rudder deflection.

detailed description

The invention will be described in further detail below in conjunction with the accompanying drawings.

In Fig. 1, the fast-loop torque command M _c obtained by designing the fast-loop control law is combined with the designed coordination factor (19) to derive the coordinated torque, and the coordinated torque is finally allocated to the rudder surface through the fast-loop rudder distribution matrix The coordination commands δ _e , δ _a , δ _r , through the coordinated deflection of the rudder surface, finally realize the coordinated control of the attitude motion of the hypersonic vehicle. Figure 2, Figure 3, and Figure 4 are simulation verifications. Through the analysis and comparison of the simulation diagrams, it can be concluded that the addition of the coordination factor can effectively improve the control effect of the controller, and has a significant effect on the suppression of coupling.

The hypersonic vehicle attitude system model is:

in Represents the derivative of Ω, Ω represents the state variable of the slow loop of the system, Ω=[α,β,μ] ^T , α,β,μ are angle of attack, angle of sideslip, and roll angle respectively; Indicates the derivative of ω, ω indicates the state variable of the fast loop of the system, ω=[p,q,r] ^T , p,q,r are the roll rate, pitch rate, and yaw rate respectively; M _c indicates the fast loop torque , M _c ＝g _fδ u, where g _fδ ∈ R ^3×3 is the fast loop control surface allocation matrix of the attitude system, u=[δ _e ,δ _a ,δ _r ] ^T , where δ _e ,δ _a ,δ _r are respectively are left elevator aileron rudder, right elevator aileron rudder and rudder; f _s =[f _α ,f _β ,f _μ ] ^T , f _f =[f _p ,f _q ,f _r ] ^T ,

Among them, M and V are the mass and velocity of the aircraft respectively; is the dynamic pressure; S, c, b are the reference area, reference length, and reference width; C _{L, α} is the lift coefficient caused by the angle of attack α; C _{Y, β} is the lateral coefficient caused by the sideslip angle β, g is the gravitational coefficient of the earth, C _{l, β} , C _{l, p} , C _lr are the lift coefficients caused by β, p, r, C _{m, α} , C _{m, q} are the basic pitching moment coefficients and the pitching force caused by q Moment coefficient, C _{D, α} is the drag coefficient caused by the angle of attack α, C _{n, β} , C _{n, p} , C _{n, r} r is the yaw moment coefficient caused by β, p, r, X _cg is the center of mass The distance to the reference moment center, Assign matrix coefficients for fast loops. I _xx is the product of inertia around the x-axis, I _yy is the product of inertia around the y-axis, I _zz is the product of inertia around the z-axis, is the derivative of I _xx , is the derivative of I _yy , is the derivative of I _zz , f _s , f _f , g _s , g _f are the state matrix of the attitude system, l _aero , ma _aero , n _aero represent the three-channel torque. It can be seen from the mathematical model of the attitude system that the attitude-motion coupling can be decomposed into attitude angle coupling, inertial coupling and rudder surface control coupling.

Attitude-angle coupling can be described as:

in Respectively expressed as the attitude angle coupling of the slow loop state variables attack angle, sideslip angle and roll angle.

Inertial coupling can be described as:

Among them, f _p , f _q , fr represent the inertial coupling caused by p, q, _r in the fast loop system.

The rudder control coupling can be described as:

Where g _l is the steering coupling of the roll channel, g _m is the steering coupling of the pitch channel, and g _n is the coupling of the yaw channel, are the rolling moment coefficients caused by the right elevator, aileron rudder and rudder, respectively, is the pitching moment coefficient caused by the right elevon rudder, The pitching moment coefficient caused by the rudder, is the yaw moment coefficient caused by the right elevon rudder, The yaw moment coefficient caused by the rudder.

G _f,δ is the distribution matrix of the rudder surface, are the lateral force coefficient caused by the left elevator, the side force coefficient caused by the right elevator aileron and the side force coefficient caused by the rudder, respectively, are the yaw moment coefficient, pitch moment coefficient, roll moment coefficient and drag coefficient caused by the left elevator, is the drag coefficient caused by the right elevator aileron, is the drag coefficient caused by the rudder, is the lift coefficient caused by the left elevator, is the lift coefficient caused by the rudder, is the yaw moment coefficient caused by the right elevator aileron, is the yaw moment coefficient caused by the rudder, Assign matrix parameters for pitch channel rudder surface, g _q,δe , g _q,δa , g _q,δr are yaw channel rudder surface assignment matrix parameters, Assign matrix parameters for the roll channel rudder surface.

In formula (7), let the side slip angle change rate be 0, that is can be set to 0, then

psina-rcosα=0 (12)

It can be drawn that r=ptanα, r=ptanα is fed back to the rudder loop, so the coordination factor is designed as

Where k ₁ >0 is a design parameter, is the first component of the rudder coordination factor.

The coupling relationship between attack angle and roll angle can be described as p-σr<p≤pcosα+rsinα<p+ρr, and β is fed back to the rudder loop, then the coordination factor is designed as:

Where |sina|≠Δ _α , k ₂ , k ₃ design parameters greater than zero, is the second component of the rudder coordination factor.

In formula (8), the coordination factor of feeding back β, r to the aileron circuit, feeding back β, q to the rudder circuit, and feeding back α, p to the elevator circuit is designed as:

Among them, λ _e , λ _a , λ _r3 are the coordination factors of corresponding state variables fed back to the left elevator, right elevon rudder and rudder, k ₄ , k ₅ , k ₆ , k ₇ , k ₈ , k ₉ are greater than zero design parameters.

For the rudder control coupling in formula (9), the coupling degree of aileron rudder and rudder control is defined as:

in,

in, is the incremental coefficient of the three-channel moment coefficient change caused by the rudder deflection, is the three-channel moment coefficient change increment coefficient caused by the right elevon rudder deflection, S _cwr , S _cw are expressed as rudder area and vertical tail area; S, L reference area and reference length; y _cwr is the distance from the rudder surface center to the longitudinal axis Distance; χ is the sweep angle; ξ is the correction factor; satisfy S _cwr ≤ S, S _cw ≤ S, y _cwr ≤ L, 0<ξ≤1, 0<cosχ≤1; n is the relative execution efficiency of the aileron; η _k , are the heel-to-tip ratio and stem-to-span ratio of the exposed wing, respectively, f the ratio of the exposed wing to the root, The rudder surface coordination strategy is designed as

δ _a = Eδ _r , 0<E≤1 (18)

E is the coupling degree of rudder and right elevon rudder. Considering equations (13), (14), and (15), the hypersonic vehicle attitude motion coordination factor is:

Among them, λ _e , λ _r , λ _a are the coordination factors of corresponding state variables fed back to the left elevator, rudder and right elevon rudder

Introducing the uncertainty vector, the attitude system model can be re-established as:

Among them, D stands for external disturbance,

θ ₁ ＝ _CL,α ,θ ₂ ＝C _Y,β ,θ ₃ ＝[C _Y,β C _L,α ] ^T ,

where C _L,α , C _Y,β , C _Y,β , C _L,α , C _n,β , C _n,p , C _n,r , C _m,α , C _m,q , C _D,α , C _l,β , C _l,r , C _l,p are the aerodynamic parameters of the aircraft.

Define Θ _Ω , Θ _ω as the bounds of uncertain parameter vectors θ _Ω , θ _ω respectively, which can be expressed as:

Among them, θ _Ωi , θ _{Ωi min} , θ _{Ωi max} are the components of the slow loop state variable, the minimum value of the component and the maximum value of the component, θ _ωi , θ _{ωi min} , θ _{ωi max} are the components of the fast loop state variable, the minimum value of the component The maximum value and component. make Expressed as an estimate of θ _Ω , Represents the estimation error, θ _ω is similar to θ _Ω , according to the projection mapping algorithm, the adaptive law of uncertain parameters is

in are fast and slow loop adaptive functions respectively, Γ>0 is a diagonal matrix, τ is an adaptive function, and the projection operator in is a projection operator, defined as

in θ _imax , θ _imin are the estimated value component of the state variable, the maximum value of the state variable component and the minimum value of the state variable component. Obviously, for the adaptive function τ

in is the adaptive estimated value, θ _min and θ _max are the estimated minimum and maximum values, For the estimation error transpose, Γ ^-1 is the inverse of the parameter diagonal matrix, Γ is the design parameter, and τ is the adaptive function.

Define the tracking error e ₁ =Ω-Ω _c , where Ω _c is the attitude angle command signal, and the sliding mode function is designed as

Wherein, K=diag{K ₁ , K ₂ , K ₃ }, K _i >0, i=1, 2, 3, which represent design parameters, and σ is a sliding mode function. Taking the first derivative of the sliding mode function gives

obtain

Where g _s , f ₁ , Ψ is the system coefficient matrix, Ψ ^T is the transpose matrix of Ψ, K is the design parameter matrix, κ ₁ >0 is the controller parameter, e ₁ is the error, The slow loop follows the derivative of the instruction, is the estimated value of the state variable of the slow loop, and its adaptive law is designed as

Where Γ ₁ ∈ R ^4×4 ,λ _Ω =diag{λ ₁ ,λ ₂ ,λ ₃ ,λ ₄ }＞0 means the design parameter matrix, is the estimation error, δ is the sliding mode function, and Ψ is the coefficient matrix of the slow loop system. Considering the Lyapunov function

where σ is the sliding mode function, σ ^T is the transpose of the planned meter function, slow loop estimation error matrix transpose, To design the inverse matrix of the parameter matrix for self-adaption, the derivative of (30) can be obtained

Combining formula (28), (29) can get

Where κ ₁ ＞0, Γ ₁ ∈R ^4×4 , λ _Ω =diag{λ ₁ ,λ ₂ ,λ ₃ ,λ ₄ }＞0 means the design parameter matrix, is the estimation error, σ is the sliding mode function, σ ^T is the transpose of the planned meter function, The slow loop estimation error matrix is transposed, g _s , f ₁ , Ψ is the system coefficient matrix, Ψ ^T is the transposed matrix of Ψ, e ₂ is the error, and g _s is the slow loop state system matrix.

Define the error e ₂ =ω-ω _c , ω _c is the command signal of the attitude angular rate, and the derivative of the error can be obtained

where Ξ ^T is the transpose of the fast loop system matrix of the system, θ _ω is the state variable of the fast loop, g _f is the system state matrix, M _c is the control torque, D is the external disturbance, is the derivative of the slow loop controller. virtual controller As a disturbance, it can be concluded that in, The attitude angle rate loop control is designed as

in is the inverse of the state matrix g _f of the fast loop system, κ ₂ is the control design parameter, is the observer output, is the transposition of the slow loop state system matrix, and other parameters are the same as above. Estimated by the observer, defining the observation error Adaptive law is

Where κ ₂ ＞0, Γ ₂ ∈R ^17×17 ＞0, λ _ω ∈R ^17×17 ＞0 are the design parameters, for the estimation error. The observer is designed as

in,

q ₁ (e), q ₂ (e),...,q _n (e) are observer parameters.

Proof: Consider the selection of the Lyapunov function as follows

Among them, e ₂ is the error, is the transposition of the error, θ _ω is the state variable of the fast loop, is the estimation error, for the transposition of is the observation error, is the observation error transpose, is the transposition matrix of the adaptive parameter, and the derivative of the Lyapunov function for time t can be obtained

in is the derivative of the slow loop Lyapunov function with respect to time t, error transpose, error derivative, for the transposition of is the fast loop adaptive function, is the transposed matrix of the observation error, The derivative of the observer output, substituting equations (32), (34), and (35) into equation (38) to get

Let L(e)=c>1, Q(e)=ce, then we can get

Among them, c is the design parameter of the observer, and other parameters are consistent with the previous ones.

From equation (40), the Lyapunov function is uniformly bounded, so the designed control law ensures that the controlled system is uniformly bounded.

The robust coordination control of attitude motion is to introduce the coordination factor (19), combine the attitude angle rate loop controller M _c and the rudder surface distribution matrix g _fδ , calculate the coordination torque, and distribute the coordination torque into rudder surface commands, which are realized by the rudder surface deflection Coordination, the robust coordination control scheme is shown in Fig. 1. In Fig. 1, the fast-loop torque command M _c obtained by designing the fast-loop control law is combined with the designed coordination factor (19) to derive the coordinated torque, and the coordinated torque is finally allocated to the rudder surface through the fast-loop rudder distribution matrix The coordination commands δ _e , δ _a , δ _r , through the coordinated deflection of the rudder surface, finally realize the coordinated control of the attitude motion of the hypersonic vehicle. Then the coordination moment can be expressed as

M _c =g _Γfδ ·δ (41)

in

in Assign matrix parameters for the pitch channel rudder surface, Assign matrix parameters for the yaw channel rudder surface, Assign matrix parameters to the roll channel rudder surface, λ _e , λ _a , λ _r are the coordination factors designed above. Then the rudder command can be obtained

δ＝g _Γfδ ^-1 ·M _c (43)

When the coordination strategy uses the coordinated deflection of the aileron rudder and the rudder to realize the yaw moment and roll moment, and the deflection of the elevator to realize the pitch moment, the rudder surface coordination can be designed as

δ _coe = δ _e , δ _coa = Eδ _r , δ _cor = δ _r (44)

In order to prove the effectiveness of the controller, we select the hypersonic vehicle attitude system model described in formula (1), the basic parameters of the model are: M ₌ 54013lb, Ma=8, V=13398ft/s, H=68898ft, T = 280000 lb. The design parameters are selected as: ki =1, _i =1...7, K ₁ =K ₂ =K ₃ =4, κ ₁ =κ ₂ =5. The initial conditions are set as: α=1°, β=5°, μ=0°, p=q=r=0°/s. The command signal of attitude angle is taken as: α _c =5°, β _c =0°, μ _c =4°. The external disturbance D is selected as D(t)=[0.5sin(1.5t) 0.5cos(2t) 0.5sin(2t)] ^T . The simulation results are shown in Figure 2, Figure 3, and Figure 4. In Figure 2-(a) when the coordination factor (14) is not added, the overshoot of α is 35%, and after adding the coordination factor (14), the overshoot is 10%. In Figure 2-(b), the overshoot of β after adding the coordination factor (14) is smaller than the overshoot without adding the coordination factor. After adding the coordination factor (15) to the fast loop, the jitter amplitudes of the tracking curves in Figure 3-(a)p, 3-(b)q and 3-(c)r are significantly reduced, and the attitude angular rate The variation range and jitter amplitude are smaller than the response curve under the action of an uncoordinated controller. Thanks to the coordinated controller, the process of attitude angular rate approaching equilibrium is smoother and faster. In Fig. 4, adding the coordination factor (18) improves the control efficiency of the rudder surface. This performance boost could improve the controllability of the attitude system and the maneuverability of the hypersonic vehicle.

Claims (5)

1. a hypersonic vehicle attitude-motion coordinated control method based on coordination factors, is characterized in that: comprise the steps: Step 1) Carry out coupling analysis for the attitude model of the hypersonic vehicle, and decompose the strong coupling problem between attitude motions into three forms: angular motion coupling, inertial coupling, and rudder surface control coupling; based on the three forms of coupling, the corresponding state variables are fed back to the corresponding rudder surface circuit design coordination factor; Step 2) Based on the time-scale separation principle, decompose the attitude system into a slow loop and a fast loop; design a slow loop robust controller and a fast loop robust controller based on the sliding mode method and projection mapping method respectively; Step 3) Combining the coordination factor with the robust controller to derive the coordination torque, using the rudder distribution matrix, the coordination torque is distributed into the rudder command, and the coordination deflection of the rudder is used to realize the coordination of the attitude movement. 2. the hypersonic vehicle attitude-motion coordinated control method based on coordination factor according to claim 1, is characterized in that, the specific process of described step 1) is following steps: Step 1-1), establishing a hypersonic vehicle attitude system mathematical model; <mfenced open = "{" close = ""><mtable><mtr><mtd><mover><mi>&amp;Omega;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><msub><mi>f</mi><mi>s</mi></msub><mo>+</mo><msub><mi>g</mi><mi>s</mi></msub><mi>&amp;omega;</mi></mtd></mtr><mtr><mtd><mover><mi>&amp;omega;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><msub><mi>f</mi><mi>f</mi></msub><mo>+</mo><msub><mi>g</mi><mi>f</mi></msub><msub><mi>M</mi><mi>c</mi></msub></mtd></mtr></mtable></mfenced> in Represents the derivative of Ω, Ω represents the state variable of the slow loop of the system, Ω=[α,β,μ] ^T , α,β,μ are angle of attack, angle of sideslip, and roll angle respectively; Indicates the derivative of ω, ω indicates the state variable of the fast loop of the system, ω=[p,q,r] ^T , p,q,r are the roll rate, pitch rate, and yaw rate respectively; M _c indicates the fast loop torque , M _c ＝g _fδ u, where g _fδ ∈ R ^3×3 is the fast loop control surface assignment matrix of the attitude system, u=[δ _e ,δ _a ,δ _r ] ^T , where δ _e ,δ _a ,δ _r are respectively are left elevator aileron rudder, right elevator aileron rudder and rudder; f _s =[f _α ,f _β ,f _μ ] ^T , f _f =[f _p ,f _q ,f _r ] ^T , <mrow><msub><mi>f</mi><mi>&amp;alpha;</mi></msub><mo>=</mo><mfrac><mn>1</mn><mrow><mi>M</mi><mi>V</mi><mi></mi><mi>cos</mi><mi>&amp;beta;</mi></mrow></mfrac><mrow><mo>(</mo><mo>-</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>L</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><mo>+</mo><mi>M</mi><mi>g</mi><mi></mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;gamma;</mi><mi>cos</mi><mi>&amp;mu;</mi><mo>)</mo></mrow></mrow> <mrow><msub><mi>f</mi><mi>&amp;beta;</mi></msub><mo>=</mo><mfrac><mn>1</mn><mrow><mi>M</mi><mi>V</mi></mrow></mfrac><mrow><mo>(</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>Y</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><mi>&amp;beta;</mi><mo>+</mo><mi>M</mi><mi>g</mi><mi></mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;gamma;</mi><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;mu;</mi><mo>)</mo></mrow></mrow> <mrow><msub><mi>f</mi><mi>&amp;mu;</mi></msub><mo>=</mo><mfrac><mn>1</mn><mrow><mi>M</mi><mi>V</mi></mrow></mfrac><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>Y</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><mi>&amp;beta;</mi><mi>tan</mi><mi>&amp;gamma;</mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;mu;</mi><mo>+</mo><mfrac><mn>1</mn><mrow><mi>M</mi><mi>V</mi></mrow></mfrac><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>L</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub></mrow> <mrow><mo>(</mo><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;gamma;</mi><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;mu;</mi><mo>+</mo><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;beta;</mi><mo>)</mo><mo>-</mo><mfrac><mi>g</mi><mi>V</mi></mfrac><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;gamma;</mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;mu;</mi><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;beta;</mi></mrow> <mrow><msub><mi>g</mi><mi>s</mi></msub><mo>=</mo><mfenced open = "[" close = "]"><mtable><mtr><mtd><mrow><mo>-</mo><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;alpha;</mi><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;beta;</mi></mrow></mtd><mtd><mn>1</mn></mtd><mtd><mrow><mo>-</mo><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;alpha;</mi><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;beta;</mi></mrow></mtd></mtr><mtr><mtd><mrow><mi>sin</mi><mi>&amp;alpha;</mi></mrow></mtd><mtd><mn>0</mn></mtd><mtd><mrow><mo>-</mo><mi>cos</mi><mi>&amp;alpha;</mi></mrow></mtd></mtr><mtr><mtd><mrow><mi>cos</mi><mi>&amp;alpha;</mi><mi>s</mi><mi>e</mi><mi>c</mi><mi>&amp;beta;</mi></mrow></mtd><mtd><mn>0</mn></mtd><mtd><mrow><mi>sin</mi>mi><mi>&amp;alpha;</mi><mi>sec</mi><mi>&amp;beta;</mi></mrow></mtd></mtr></mtable></mfenced></mrow> <mrow><msub><mi>f</mi><mi>p</mi></msub><mo>=</mo><mfrac><mn>1</mn><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mfrac><mrow><mo>(</mo><msub><mi>l</mi><mrow><mi>a</mi><mi>e</mi><mi>r</mi><mi>o</mi></mrow></msub><mo>+</mo><mo>(</mo><mrow><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mrow><mo>)</mo><mi>q</mi><mi>r</mi><mo>-</mo><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>x</mi><mi>x</mi></mrow></msub><mi>p</mi><mo>)</mo></mrow></mrow> <mrow><msub><mi>f</mi><mi>q</mi></msub><mo>=</mo><mfrac><mn>1</mn><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac><mrow><mo>(</mo><msub><mi>m</mi><mrow><mi>a</mi><mi>e</mi><mi>r</mi><mi>o</mi></mrow></msub><mo>+</mo><mo>(</mo><mrow><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mrow><mo>)</mo><mi>p</mi><mi>r</mi><mo>-</mo><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>y</mi><mi>y</mi></mrow></msub><mi>q</mi><mo>)</mo></mrow></mrow> <mrow><msub><mi>f</mi><mi>r</mi></msub><mo>=</mo><mfrac><mn>1</mn><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac><mrow><mo>(</mo><msub><mi>n</mi><mrow><mi>a</mi><mi>e</mi><mi>r</mi><mi>o</mi></mrow></msub><mo>+</mo><mo>(</mo><mrow><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mrow><mo>)</mo><mi>p</mi><mi>q</mi><mo>-</mo><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>z</mi><mi>z</mi></mrow></msub><mi>r</mi><mo>)</mo></mrow></mrow> <mrow><msub><mi>g</mi><mi>f</mi></msub><mo>=</mo><mfenced open = "[" close = "]"><mtable><mtr><mtd><msubsup><mi>g</mi><mi>l</mi><mi>p</mi></msubsup></mtd><mtd><mn>0</mn></mtd><mtd><mn>0</mn></mtd></mtr><mtr><mtd><mn>0</mn></mtd><mtd><msubsup><mi>g</mi><mi>m</mi><mi>q</mi></msubsup></mtd><mtd><mn>0</mn></mtd></mtr><mtr><mtd><mn>0</mn></mtd><mtd><mn>0</mn></mtd><mtd><msubsup><mi>g</mi><mi>n</mi><mi>r</mi></msubsup></mtd></mtr></mtable></mfenced></mrow> <mrow><msub><mi>l</mi><mrow><mi>a</mi><mi>e</mi><mi>r</mi><mi>o</mi></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>b</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><mi>&amp;beta;</mi><mo>+</mo><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><mi>p</mi></mrow></msub><mi>p</mi><mi>b</mi><mo>/</mo><mn>2</mn><mi>V</mi><mo>+</mo><msub><mi>C</mi><mrow><mi>l</mi><mo>.</mo><mi>r</mi></mrow></msub><mi>r</mi><mi>b</mi><mo>/</mo><mn>2</mn><mi>V</mi><mo>)</mo></mrow><mo>,</mo></mrow> 1 <mrow><mtable><mtr><mtd><mrow><msub><mi>m</mi><mrow><mi>a</mi><mi>e</mi><mi>r</mi><mi>o</mi></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>c</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>m</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><mo>+</mo><msub><mi>C</mi><mrow><mi>m</mi><mo>,</mo><mi>q</mi></mrow></msub><mi>q</mi><mi>c</mi><mo>/</mo><mn>2</mn><mi>V</mi><mo>)</mo></mrow><mo>+</mo></mrow></mtd></mtr><mtr><mtd><mrow><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>D</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><mi>sin</mi><mi>&amp;alpha;</mi><mo>+</mo><msub><mi>C</mi><mrow><mi>L</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><mi>cos</mi><mi>&amp;alpha;</mi><mo>)</mo></mrow></mrow></mtd></mtr></mtable><mo>,</mo></mrow> <mfenced open=""close=""><mtable><mtr><mtd><mrow><msub><mi>n</mi><mrow><mi>a</mi><mi>e</mi><mi>r</mi><mi>o</mi></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>b</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><mi>&amp;beta;</mi><mo>+</mo><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><mi>p</mi></mrow></msub><mi>p</mi><mi>b</mi><mo>/</mo><mn>2</mn><mi>V</mi><mo>+</mo><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><mi>r</mi></mrow></msub><mi>r</mi><mi>b</mi><mo>/</mo><mn>2</mn><mi>V</mi><mo>)</mo></mrow><mo>+</mo></mrow></mtd></mtr><mtr><mtd><mrow><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>Y</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><mi>&amp;beta;</mi></mrow></mtd></mtr></mtable></mfenced> Among them, M and V are the mass and velocity of the aircraft respectively; is the dynamic pressure; S, c, b are the reference area, reference length, and reference width respectively; C _{L, α} is the lift coefficient caused by the angle of attack α, C _{Y, β} is the lateral coefficient caused by the sideslip angle β, g is the gravitational coefficient of the earth, C _{l, β} , C _{l, p} , C _lr are the lift coefficients caused by β, p, r, C _{m, α} , C _{m, q} are the basic pitching moment coefficients and the pitching force caused by q Moment coefficient, C _{D, α} is the drag coefficient caused by the angle of attack α, C _{n, β} , C _{n, p} , C _{n, r} r is the yaw moment coefficient caused by β, p, r, X _cg is the center of mass The distance to the reference moment center, Assign matrix coefficients for the fast loop; I _xx is the product of inertia around the x-axis, I _yy is the product of inertia around the y-axis, I _zz is the product of inertia around the z-axis, is the derivative of I _xx , is the derivative of I _yy , is the derivative of I _zz , f _s , f _f , g _s , g _f are the attitude system state matrix, l _aero , m _aero , n _aero represent the three-channel torque; Step 1-2) Carry out coupling analysis to the established attitude system model, decompose the coupling of attitude attitude into attitude angle motion coupling, inertial coupling, rudder surface control coupling, 1) Attitude Angle Coupling Model: In the attitude system model established in step 1-1), the coupling relation of the attitude angle is described as: <mrow><mover><mi>&amp;alpha;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><mo>-</mo><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;beta;</mi><mrow><mo>(</mo><mi>p</mi><mi></mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;alpha;</mi><mo>+</mi>mo><mi>r</mi><mi></mi><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;alpha;</mi><mo>)</mo></mrow></mrow> <mrow><mover><mi>&amp;beta;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><mi>p</mi><mi></mi><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;alpha;</mi><mo>-</mo><mi>r</mi><mi></mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;alpha;</mi></mrow> <mrow><mover><mi>&amp;mu;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><mi>sec</mi><mi>&amp;beta;</mi><mrow><mo>(</mo><mi>p</mi><mi></mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;alpha;</mi><mo>+</mo><mi>r</mi><mi></mi><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;alpha;</mi><mo>)</mo></mrow></mrow> in Respectively expressed as slow loop state variables attack angle, sideslip angle and attitude angle coupling of roll angle; 2) Inertial coupling In the attitude system model established in step 1-1), the inertial coupling is described as: <mrow><msub><mi>f</mi><mi>p</mi></msub><mo>=</mo><mrow><mo>(</mo><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub><mo>)</mo></mrow><mi>q</mi><mi>r</mi><mo>-</mo><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>x</mi><mi>x</mi></mrow></msub><mi>p</mi></mrow> <mrow><msub><mi>f</mi><mi>q</mi></msub><mo>=</mo><mrow><mo>(</mo><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub><mo>)</mo></mrow><mi>p</mi><mi>r</mi><mo>-</mo><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>y</mi><mi>y</mi></mrow></msub><mi>q</mi></mrow> <mrow><msub><mi>f</mi><mi>r</mi></msub><mo>=</mo><mrow><mo>(</mo><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub><mo>)</mo></mrow><mi>p</mi><mi>q</mi><mo>-</mo><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>x</mi><mi>z</mi></mrow></msub><mi>p</mi></mrow> where f _p , f _q , f _r represent the inertial coupling caused by p, q, r in the fast loop system; 3) Rudder control coupling In the attitude system model established in step 1-1), the control coupling of the rudder surface is described as: <mrow><msub><mi>g</mi><mi>l</mi></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>b</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>+</mo><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>)</mo></mrow></mrow> <mrow><msub><mi>g</mi><mi>m</mi></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>c</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>m</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>+</mo><msub><mi>C</mi><mrow><mi>m</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>)</mo></mrow></mrow> <mrow><msub><mi>g</mi><mi>n</mi></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>b</mi><mrow><mo>(</mo><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>+</mo><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>)</mo></mrow></mrow> Where g _l is the steering coupling of the roll channel, g _m is the steering coupling of the pitch channel, and g _n is the coupling of the yaw channel, are the rolling moment coefficients caused by the right elevator, aileron rudder and rudder, respectively, is the pitching moment coefficient caused by the right elevon rudder, The pitching moment coefficient caused by the rudder, is the yaw moment coefficient caused by the right elevon rudder, yaw moment coefficient caused by rudder; <mrow><msub><mi>G</mi><mrow><mi>f</mi><mo>,</mo><mi>&amp;delta;</mi></mrow></msub><mo>=</mo><mfenced open = "[" close = "]"><mtable><mtr><mtd><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub></mtd></mtr><mtr><mtd><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub></mtd></mtr><mtr><mtd><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub></mtd></mtr></mtable></mfenced></mrow> <mrow><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SbC</mi><mrow><mi>l</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub></mrow> <mrow><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SbC</mi><mrow><mi>l</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mrow> <mrow><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SbC</mi><mrow><mi>l</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub></mrow> <mfenced open = "" close = ""><mtable><mtr><mtd><mrow><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>ScC</mi><mrow><mi>m</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mo>+</mo><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mo>(</mo><msub><mi>C</mi><mrow><mi>D</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mi>sin</mi><mi>&amp;alpha;</mi></mrow></mtd></mtr><mtr><mtd><mrow><mo>+</mo><msub><mi>C</mi><mrow><mi>L</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mi>cos</mi><mi>&amp;alpha;</mi><mo>)</mo></mrow></mtd></mtr></mtable></mfenced> <mfenced open = "" close = ""><mtable><mtr><mtd><mrow><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>ScC</mi><mrow><mi>m</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>+</mo><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mo>(</mo><msub><mi>C</mi><mrow><mi>D</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mi>sin</mi><mi>&amp;alpha;</mi></mrow></mtd></mtr><mtr><mtd><mrow><mo>+</mo><msub><mi>C</mi><mrow><mi>L</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mi>cos</mi><mi>&amp;alpha;</mi><mo>)</mo></mrow></mtd></mtr></mtable></mfenced> <mrow><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>ScC</mi><mrow><mi>m</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>+</mo><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>D</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mi>sin</mi><mi>&amp;alpha;</mi></mrow> <mrow><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SbC</mi><mrow><mi>n</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mo>+</mo><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>Y</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub></mrow> <mrow><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SbC</mi><mrow><mi>n</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub><mo>+</mo><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>Y</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mrow> <mrow><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>=</mo><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SbC</mi><mrow><mi>n</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>+</mo><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msub><mi>SC</mi><mrow><mi>Y</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub></mrow> G _f,δ is the distribution matrix of the rudder surface, are the lateral force coefficient caused by the left elevator, the side force coefficient caused by the right elevator aileron and the side force coefficient caused by the rudder, respectively, are the yaw moment coefficient, pitch moment coefficient, roll moment coefficient and drag coefficient caused by the left elevator, is the drag coefficient caused by the right elevator aileron, is the drag coefficient caused by the rudder, is the lift coefficient caused by the left elevator, is the lift coefficient caused by the rudder, is the yaw moment coefficient caused by the right elevator aileron, is the yaw moment coefficient caused by the rudder, Assign matrix parameters for pitch channel rudder surface, g _q,δe , g _q,δa , g _q,δr are yaw channel rudder surface assignment matrix parameters, Assign matrix parameters for the roll channel rudder surface; Step 1-3) the design of coordination factor; For 1) attitude angle coupling in steps 1-2), Set to 0, expressed as psina-rcosα=0, get r=ptanα, r=ptanα is fed back to the rudder loop, so the coordination factor is designed as Where k ₁ >0 is a design parameter, is the first component of the rudder coordination factor; In steps 1-2), 1) the coupling relationship between attack angle and roll angle is described as p-σr<p≤pcosα+rsinα<p+ρr, and the adjustment factor for feeding β to the rudder loop is designed as: <mrow><msub><mi>&amp;lambda;</mi><msub><mi>r</mi><mn>2</mn></msub></msub><mo>=</mo><msub><mi>k</mi><mn>2</mn></msub><mi>&amp;beta;</mi><mo>+</mo><msub><mi>k</mi><mn>3</mn></msub><mfrac><mrow><mi>p</mi><mrow><mo>(</mo><mn>1</mn><mo>-</mo><mi>cos</mi><mi>&amp;alpha;</mi><mo>)</mo></mrow></mrow><mrow><mi>sin</mi><mi>&amp;alpha;</mi><mo>+</mo><msub><mi>&amp;Delta;</mi><mi>&amp;alpha;</mi></mi></mi>msub></mrow></mfrac></mrow> Where |sina|≠Δ _α , k ₂ , k ₃ design parameters greater than zero, is the second component of the rudder coordination factor; For 2) inertial coupling in step 1-2), feed β,r to the aileron loop, β,q to the rudder loop, and α,p to the elevator loop to increase the damping torque and stability torque; coordination The factorial design is: <mfenced open = "{" close = ""><mtable><mtr><mtd><msub><mi>&amp;lambda;</mi><mi>e</mi></msub><mo>=</mo><msub><mi>k</mi><mn>4</mn></msub><mi>&amp;alpha;</mi><mo>+</mo><msub><mi>k</mi><mn>5</mn></msub><mi>p</mi></mtd></mtr><mtr><mtd><msub><mi>&amp;lambda;</mi><mi>a</mi></msub><mo>=</mo><msub><mi>k</mi><mn>6</mn></msub><mi>&amp;beta;</mi><mo>+</mo><msub><mi>k</mi><mn>7</mn></msub><mi>r</mi></mtd></mtr><mtr><mtd><msub><mi>&amp;lambda;</mi><msub><mi>r</mi><mn>3</mn></msub></msub><mo>=</mo><msub><mi>k</mi><mn>8</mn></msub><mi>&amp;beta;</mi><mo>+</mo><msub><mi>k</mi><mn>9</mn></msub><mi>q</mi></mtd></mtr></mtable></mfenced> 3 where λ _e , λ _a , are respectively the coordination factors corresponding to the state variables fed back to the left elevator, right elevon rudder and rudder, k ₄ , k ₅ , k ₆ , k ₇ , k ₈ , k ₉ are design parameters greater than zero; For 3) rudder control coupling in steps 1-2), define the coupling degree of aileron rudder and rudder control as: <mrow><mi>E</mi><mo>=</mo><mfrac><msubsup><mi>C</mi><mi>x</mi><msub><mi>&amp;delta;</mi><mi>r</mi></msub></msubsup><msubsup><mi>C</mi><mi>x</mi><msub><mi>&amp;delta;</mi><mi>a</mi></msub></msubsup></mfrac><mo>&amp;times;</mo><mn>100</mn><mi>%</mi></mrow> in, <mrow><msubsup><mi>C</mi><mi>x</mi><msub><mi>&amp;delta;</mi><mi>r</mi></msub></msubsup><mo>=</mo><mfrac><mrow><msub><mi>S</mi><mrow><mi>c</mi><mi>w</mi><mi>r</mi></mrow></msub><msub><mi>S</mi><mrow><mi>c</mi><mi>w</mi></mrow></msub></mrow><msup><mi>S</mi><mn>2</mn></msup></mfrac><mfrac><msub><mi>y</mi><mrow><mi>c</mi><mi>w</mi><mi>r</mi></mrow></msub><mi>L</mi></mfrac><mi>&amp;xi;</mi><mi>cos</mi><mi>&amp;chi;</mi></mrow> <mrow><msubsup><mi>C</mi><mi>x</mi><msub><mi>&amp;delta;</mi><mi>a</mi></msub></msubsup><mo>=</mo><mfrac><mn>1</mn><mn>2</mn></mfrac><mo>&amp;lsqb;</mo><mi>n</mi><mfrac><mrow><mo>(</mo><msub><mi>&amp;eta;</mi><mi>k</mi></msub><mo>+</mo><mn>1</mn><mo>)</mo><msup><mrow><mo>(</mo><mn>1</mn><mo>-</mo><mover><mi>D</mi><mo>&amp;OverBar;</mo></mover><mo>)</mo></mrow><mn>2</mn></msup></mrow><mrow><msub><mi>&amp;eta;</mi><mi>k</mi></msub><mo>+</mo><mn>1</mn><mo>-</mo><mn>2</mn><mover><mi>D</mi><mo>&amp;OverBar;</mo></mover></mrow></mfrac><mo>&amp;lsqb;</mo><mover><mi>D</mi><mo>&amp;OverBar;</mo></mover><mo>+</mo><mrow><mo>(</mo><mn>1</mn><mo>-</mo><mover><mi>D</mi><mo>&amp;OverBar;</mo></mover><mo>)</mo></mrow><mi>f</mi><mo>&amp;rsqb;</mo><mo>&amp;rsqb;</mo></mrow> in, is the incremental coefficient of the three-channel moment coefficient change caused by the rudder deflection, is the three-channel moment coefficient change increment coefficient caused by the right _elevon rudder deflection, S _cwr , S _cw are expressed as rudder area and vertical tail area; S, L reference area and reference length; Distance; χ is the sweep angle; ξ is the correction factor; satisfy S _cwr ≤ S, S _cw ≤ S, y _cwr ≤ L, 0<ξ≤1, 0<cosχ≤1; n is the relative execution efficiency of the aileron; η _k , are the heel-to-tip ratio and stem-to-span ratio of the exposed wing, respectively; f is the ratio of the exposed wing to the root; the coordination strategy of the rudder surface is designed as δ _a = Eδ _r , 0<E≤1, and E is the coupling degree of the rudder and the right elevon rudder. 3. the hypersonic vehicle attitude-motion coordinated control method based on coordination factor according to claim 1, is characterized in that, described step 2) the concrete process following steps: Step 2-1) Introduce uncertain parameter vector By introducing an uncertain parameter vector, the system model in step 1-1) is expressed as <mrow><mover><mi>&amp;Omega;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><msub><mi>f</mi><mn>1</mn></msub><mo>+</mo><msup><mi>&amp;Psi;</mi><mi>T</mi></msup><msub><mi>&amp;theta;</mi><mi>&amp;Omega;</mi></msub><mo>+</mo><msub><mi>g</mi><mi>s</mi></msub><mi>&amp;omega;</mi></mrow> <mrow><mover><mi>&amp;omega;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><msup><mi>&amp;Xi;</mi><mi>T</mi></msup><msub><mi>&amp;theta;</mi><mi>&amp;omega;</mi></msub><mo>+</mo><msub><mi>g</mi><mi>f</mi></msub><msub><mi>M</mi><mi>c</mi></msub><mo>+</mo><mi>D</mi></mrow> Among them, D stands for external disturbance, <mrow><msub><mi>&amp;theta;</mi><mi>&amp;Omega;</mi></msub><mo>=</mo><msubsup><mfenced open = "[" close = "]"><mtable><mtr><mtd><msubsup><mi>&amp;theta;</mi><mn>1</mn><mi>T</mi></msubsup></mtd><mtd><msubsup><mi>&amp;theta;</mi><mn>2</mn><mi>T</mi></msubsup></mtd><mtd><msubsup><mi>&amp;theta;</mi><mn>3</mn><mi>T</mi></msubsup></mtd></mtr></mtable></mfenced><mrow><mn>4</mn><mo>&amp;times;</mo><mn>1</mn></mrow><mi>T</mi></msubsup><mo>,</mo><msub><mi>&amp;theta;</mi><mi>&amp;omega;</mi></msub><mo>=</mo><msubsup><mfenced open = "[" close = "]"><mtable><mtr><mtd><msubsup><mi>&amp;theta;</mi><mn>4</mn><mi>T</mi></msubsup></mtd><mtd><msubsup><mi>&amp;theta;</mi><mn>5</mn><mi>T</mi></msubsup></mtd><mtd><msubsup><mi>&amp;theta;</mi><mn>6</mn><mi>T</mi></msubsup></mtd></mtr></mtable></mfenced><mrow><mn>17</mn><mo>&amp;times;</mo><mn>1</mn></mrow><mi>T</mi></msubsup></mrow> θ ₁ ＝ _CL,α ,θ ₂ ＝C _Y,β ,θ ₃ ＝[C _Y,β C _L,α ] ^T , <mrow><msub><mi>&amp;theta;</mi><mn>4</mn></msub><mo>=</mo><msup><mfenced open = "[" close = "]"><mtable><mtr><mtd><mfrac><mrow><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mrow><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>z</mi><mi>z</mi></mrow></msub><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><mi>p</mi></mrow></msub><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>n</mi><mo>,</mo><mi>r</mi></mrow></msub><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>Y</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mfrac></mtd></mtr></mtable></mfenced><mi>T</mi></msup><mo>,</mo></mrow> <mrow><msub><mi>&amp;theta;</mi><mn>5</mn></msub><mo>=</mo><msup><mfenced open = "[" close = "]"><mtable><mtr><mtd><mfrac><mrow><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mrow><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>y</mi><mi>y</mi></mrow></msub><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>m</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>m</mi><mo>,</mo><mi>q</mi></mrow></msub><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>D</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>L</mi><mo>,</mo><mi>&amp;alpha;</mi></mrow></msub><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub></mfrac></mtd></mtr></mtable></mfenced><mi>T</mi></msup><mo>,</mo></mrow> <mrow><msub><mi>&amp;theta;</mi><mn>6</mn></msub><mo>=</mo><msup><mfenced open = "[" close = "]"><mtable><mtr><mtd><mfrac><mrow><msub><mi>I</mi><mrow><mi>y</mi><mi>y</mi></mrow></msub><mo>-</mo><msub><mi>I</mi><mrow><mi>z</mi><mi>z</mi></mrow></msub></mrow><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mover><mi>I</mi><mo>&amp;CenterDot;</mo></mover><mrow><mi>x</mi><mi>x</mi></mrow></msub><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><mi>&amp;beta;</mi></mrow></msub><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><mi>p</mi></mrow></msub><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mfrac></mtd><mtd><mfrac><msub><mi>C</mi><mrow><mi>l</mi><mo>,</mo><mi>r</mi></mrow></msub><msub><mi>I</mi><mrow><mi>x</mi><mi>x</mi></mrow></msub></mfrac></mtd></mtr></mtable></mfenced><mi>T</mi></msup><mo>,</mo></mrow> <mrow><mi>&amp;Psi;</mi><mo>=</mo><msub><mfenced open = "[" close = "]"><mtable><mtr><mtd><msub><mi>&amp;Psi;</mi><mn>1</mn></msub></mtd><mtd><mn>0</mn></mtd><mtd><mn>0</mn></mtd></mtr><mtr><mtd><mn>0</mn></mtd><mtd><msub><mi>&amp;psi;</mi><mn>2</mn></msub></mtd><mtd><mn>0</mn></mtd></mtr><mtr><mtd><msub><mn>0</mn><mrow><mn>2</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd><mtd><msub><mn>0</mn><mrow><mn>2</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd><mtd><msub><mi>&amp;Psi;</mi><mn>3</mn></msub></mtd></mtr></mtable></mfenced><mrow><mn>4</mn><mo>&amp;times;</mo><mn>3</mn></mrow></msub><mo>,</mo><mi>&amp;Xi;</mi><mo>=</mo><msub><mfenced open = "[" close = "]"><mtable><mtr><mtd><msub><mi>&amp;Xi;</mi><mn>1</mn></msub></mtd><mtd><msub><mn>0</mn><mrow><mn>5</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd><mtd><msub><mn>0</mn><mrow><mn>5</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd></mtr><mtr><mtd><msub><mn>0</mn><mrow><mn>6</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd><mtd><msub><mi>&amp;Xi;</mi><mn>2</mn></msub></mtd><mtd><msub><mn>0</mn><mrow><mn>6</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd></mtr><mtr><mtd><msub><mn>0</mn><mrow><mn>6</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd><mtd><msub><mn>0</mn><mrow><mn>6</mn><mo>&amp;times;</mo><mn>1</mn></mrow></msub></mtd><mtd><msub><mi>&amp;Xi;</mi><mn>3</mn></msub></mtd></mtr></mtable></mfenced><mrow><mn>17</mn><mo>&amp;times;</mo><mn>3</mn></mrow></msub></mrow> <mrow><msub><mi>&amp;Psi;</mi><mn>1</mn></msub><mrow><mo>(</mo><mi>x</mi><mo>)</mo></mrow><mo>=</mo><mo>-</mo><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi></mrow><mrow><mi>M</mi><mi>V</mi><mi></mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;beta;</mi></mrow></mfrac><mo>,</mo><msub><mi>&amp;Psi;</mi><mn>2</mn></msub><mo>=</mo><mo>-</mo><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>&amp;beta;</mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;beta;</mi></mrow><mrow><mi>M</mi><mi>V</mi></mrow></mfrac><mo>,</mo></mrow> 4 <mrow><msub><mi>&amp;Psi;</mi><mn>3</mn></msub><mo>=</mo><msup><mfenced open = "[" close = "]"><mtable><mtr><mtd><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>&amp;beta;</mi><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;gamma;</mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;mu;</mi><mi>c</mi><mi>o</mi><mi>s</mi><mi>&amp;beta;</mi></mrow><mrow><mi>M</mi><mi>V</mi></mrow></mfrac></mtd><mtd><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mrow><mo>(</mo><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;gamma;</mi><mi>s</mi><mi>i</mi><mi>n</mi><mi>&amp;mu;</mi><mo>+</mo><mi>t</mi><mi>a</mi><mi>n</mi><mi>&amp;beta;</mi><mo>)</mo></mrow></mrow><mrow><mi>M</mi><mi>V</mi></mrow></mfrac></mtd></mtr></mtable></mfenced><mi>T</mi></msup></mrow> <mrow><msub><mi>&amp;Xi;</mi><mn>1</mn></msub><mrow><mo>(</mo><mi>x</mi><mo>)</mo></mrow><mo>=</mo><msup><mfenced open = "[" close = "]"><mtable><mtr><mtd><mrow><mi>q</mi><mi>r</mi></mrow></mtd><mtd><mrow><mo>-</mo><mi>p</mi></mrow></mtd><mtd><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>b</mi><mi>&amp;beta;</mi></mrow></mtd><mtd><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msup><mi>Sb</mi><mn>2</mn></msup><mi>p</mi></mrow><mrow><mn>2</mn><mi>V</mi></mrow></mfrac></mtd><mtd><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msup><mi>Sb</mi><mn>2</mn></msup><mi>r</mi></mrow><mrow><mn>2</mn><mi>V</mi></mrow></mfrac></mtd></mtr></mtable></mfenced><mi>T</mi></msup><mo>,</mo></mrow> <mfenced open = "" close = ""><mtable><mtr><mtd><mrow><msub><mi>&amp;Xi;</mi><mn>2</mn></msub><mrow><mo>(</mo><mi>x</mi><mo>)</mo></mrow><mo>=</mo><mo>&amp;lsqb;</mo><mtable><mtr><mtd><mrow><mi>p</mi><mi>r</mi></mrow></mtd><mtd><mrow><mo>-</mo><mi>q</mi></mrow></mtd><mtd><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>c</mi></mrow></mtd></mtr></mtable></mrow></mtd></mtr><mtr><mtd><mtable><mtr><mtd><mfrac><mrow><msup><mi>qSc</mi><mn>2</mn></msup><mi>q</mi></mrow><mrow><mn>2</mn><mi>V</mi></mrow></mfrac></mtd><mtd><mrow><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi></mi><mi>sin</mi><mi>&amp;alpha;</mi></mrow></mtd><mtd><mrow><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi></mi><mi>cos</mi><mi>&amp;alpha;</mi><msup><mo>&amp;rsqb;</mo><mi>T</mi></msup></mrow></mtd></mtr></mtable></mtd></mtr></mtable></mfenced> <mrow><msub><mi>&amp;Xi;</mi><mn>3</mn></msub><mrow><mo>(</mo><mi>x</mi><mo>)</mo></mrow><mo>=</mo><msup><mfenced open = "[" close = "]"><mtable><mtr><mtd><mrow><mi>p</mi><mi>q</mi></mrow></mtd><mtd><mrow><mo>-</mo><mi>r</mi></mrow></mtd><mtd><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>b</mi><mi>&amp;beta;</mi></mrow></mtd><mtd><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msup><mi>Sb</mi><mn>2</mn></msup><mi>p</mi></mrow><mrow><mn>2</mn><mi>V</mi></mrow></mfrac></mtd><mtd><mfrac><mrow><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><msup><mi>Sb</mi><mn>2</mn></msup><mi>r</mi></mrow><mrow><mn>2</mn><mi>V</mi></mrow></mfrac></mtd><mtd><mrow><msub><mi>X</mi><mrow><mi>c</mi><mi>g</mi></mrow></msub><mover><mi>q</mi><mo>&amp;OverBar;</mo></mover><mi>S</mi><mi>&amp;beta;</mi></mrow></mtd></mtr></mtable></mfenced><mi>T</mi></msup></mrow> <mrow><msub><mi>f</mi><mn>1</mn></msub><mo>=</mo><mfenced open = "[" close = "]"><mtable><mtr><mtd><mfrac><mrow><mi>g</mi><mi></mi><mi>cos</mi><mi>&amp;gamma;</mi><mi>cos</mi><mi></mi><mi>u</mi></mrow><mrow><mi>V</mi><mi></mi><mi>cos</mi><mi>&amp;beta;</mi></mrow></mfrac></mtd></mtr><mtr><mtd><mfrac><mrow><mi>g</mi><mi></mi><mi>cos</mi><mi>&amp;gamma;</mi><mi>sin</mi><mi></mi><mi>u</mi></mrow><mi>V</mi></mfrac></mtd></mtr><mtr><mtd><mrow><mo>-</mo><mfrac><mrow><mi>g</mi><mi></mi><mi>cos</mi><mi>&amp;gamma;</mi><mi>cos</mi><mi></mi><mi>u</mi><mi></mi><mi>tan</mi><mi>&amp;beta;</mi></mrow><mi>V</mi></mfrac></mrow></mtd></mtr></mtable></mfenced></mrow> where C _L,α , C _Y,β , C _Y,β , C _L,α , C _n,β , C _n,p , C _n,r , C _m,α , C _m,q , C _D,α , C _l,β , C _l,r , C _l,p are the aerodynamic parameters of the aircraft; Step 2-2) Projection Mapping Algorithm Θ _Ω , Θ _ω are denoted as the bounds of uncertain parameter vectors θ _Ω , θ _ω respectively, expressed as: Θ _Ω ＝{θ _Ω ∈ R ⁴ |θ _{Ωi min ≤θ Ωi ≤θ Ωi max} ,i=1,…,4}, Θ _ω ＝{θ _ω ∈ R ¹⁷ |θ _{ωimin ≤θ ωi ≤θ ωimax} ,i=1,…,17} Among them, θ _Ωi , θ _{Ωi min} , θ _{Ωi max} are the components of the slow loop state variable, the minimum value of the component and the maximum value of the component, θ _ωi , θ _{ωi min} , θ _{ωi max} are the components of the fast loop state variable, the minimum value of the component The maximum value and component; let Expressed as an estimate of θ _Ω , Represents the estimation error, θ _ω is similar to θ _Ω , according to the projection mapping algorithm, the adaptive law of uncertain parameters is <mrow><msub><mover><mover><mi>&amp;theta;</mi><mo>^</mo></mover><mo>&amp;CenterDot;</mo></mover><mi>&amp;Omega;</mi></msub><mo>=</mo><msub><mi>Proj</mi><mover><mi>&amp;theta;</mi><mo>^</mo></mover></msub><mrow><mo>(</mo><mi>&amp;Gamma;</mi><mi>&amp;tau;</mi><mo>)</mo></mrow></mrow> <mrow><msub><mover><mover><mi>&amp;theta;</mi><mo>^</mo></mover><mo>&amp;CenterDot;</mo></mover><mi>&amp;omega;</mi></msub><mo>=</mo><msub><mi>Proj</mi><mover><mi>&amp;theta;</mi><mo>^</mo></mover></msub><mrow><mo>(</mo><mi>&amp;Gamma;</mi><mi>&amp;tau;</mi><mo>)</mo></mrow></mrow> in are fast and slow loop adaptive functions respectively, Γ>0 is a diagonal matrix, τ is an adaptive function, and the projection operator defined as in θ _imax , θ _imin is the estimated value component of the state variable, the maximum value of the state variable component and the minimum value of the state variable component Obviously, for the adaptive function τ, <mfenced open = "" close = ""><mtable><mtr><mtd><msub><mi>P</mi><mn>1</mn></msub></mtd><mtd><mrow><mover><mi>&amp;theta;</mi><mo>^</mo></mover><mo>&amp;Subset;</mo><mi>&amp;Theta;</mi><mo>=</mo><mo>{</mo><mover><mi>&amp;theta;</mi><mo>^</mo></mover><mo>|</mo><msub><mi>&amp;theta;</mi><mrow><mi>m</mi><mi>i</mi><mi>n</mi></mrow></msub><mo>&amp;le;</mo><mover><mi>&amp;theta;</mi><mo>^</mo></mover><mo>&amp;le;</mo><msub><mi>&amp;theta;</mi><mrow><mi>m</mi><mi>a</mi><mi>x</mi></mrow></msub><mo>}</mo></mrow></mtd></mtr></mtable></mfenced> <mrow><mtable><mtr><mtd><msub><mi>P</mi><mn>2</mn></msub></mtd><mtd><mrow><msup><mover><mi>&amp;theta;</mi><mo>~</mo></mover><mi>T</mi></msup><mrow><mo>(</mo><msup><mi>&amp;Gamma;</mi><mrow><mo>-</mo><mn>1</mn></mrow></msup><msub><mi>Proj</mi><mover><mi>&amp;theta;</mi><mo>^</mo></mover></msub><mo>(</mo><mrow><mi>&amp;Gamma;</mi><mi>&amp;tau;</mi></mrow><mo>)</mo><mo>-</mo><mi>&amp;tau;</mi><mo>)</mo></mrow><mo>&amp;le;</mo><mn>0</mn><mo>,</mo><mo>&amp;ForAll;</mo><mi>&amp;tau;</mi></mrow></mtd></mtr></mtable><mo>.</mo></mrow> in is the adaptive estimated value, θ _min and θ _max are the estimated minimum and maximum values, For the estimated error transposition, Γ ^-1 is the inverse of the parameter diagonal matrix, and Γ is the design parameter; Step 2-3) Attitude Angle Robust Controller Design Define the tracking error e ₁ =Ω-Ω _c , the sliding mode function is designed as <mrow><mi>&amp;sigma;</mi><mo>=</mo><msub><mi>e</mi><mn>1</mn></msub><mo>+</mo><mi>K</mi><msubsup><mo>&amp;Integral;</mo><mn>0</mn><mi>t</mi></msubsup><msub><mi>e</mi><mn>1</mn></msub><mi>d</mi><mi>t</mi></mrow> Among them, K=diag{K ₁ , K ₂ ,K ₃ }, K _i >0, i=1, 2, 3, which represent the design parameters, σ is the sliding mode function, and the first order derivative of the sliding mode function is obtained <mfenced open = "" close = ""><mtable><mtr><mtd><mrow><mover><mi>&amp;sigma;</mi><mo>&amp;CenterDot;</mo></mover><mo>=</mo><msub><mover><mi>e</mi><mo>&amp;CenterDot;</mo></mover><mn>1</mn></msub><mo>+</mo><msub><mi>Ke</mi><mn>1</mn></msub></mrow></mtd></mtr><mtr><mtd><mrow><mo>=</mo><msub><mi>f</mi><mn>1</mn></msub><mo>+</mo><msup><mi>&amp;Psi;</mi><mi>T</mi></msup><msub><mi>&amp;theta;</mi><mi>&amp;Omega;</mi></msub><mo>+</mo><msub><mi>g</mi><mi>s</mi></msub><mi>&amp;omega;</mi><mo>-</mo><msub><mover><mi>&amp;Omega;</mi><mo>&amp;CenterDot;</mo></mover><mi>c</mi></msub><mo>+</mo><msub><mi>Ke</mi><mn>1</mn></msub></mrow></mtd></mtr><mtr><mtd><mrow><mo>=</mo><msub><mi>f</mi><mn>1</mn></msub><mo>+</mo><msup><mi>&amp;Psi;</mi><mi>T</mi></msup><msub><mi>&amp;theta;</mi><mi>&amp;Omega;</mi></msub><mo>+</mo><msub><mi>g</mi><mi>s</mi></msub><mrow><mo>(</mo><msub><mi>e</mi><mn>2</mn></msub><mo>+</mo><msub><mi>&amp;omega;</mi><mi>c</mi></msub><mo>)</mo></mrow><mo>-</mo><msub><mover><mi>&amp;Omega;</mi><mo>&amp;CenterDot;</mo></mover><mi>c</mi></msub><mo>+</mo><msub><mi>Ke</mi><mn>1</mn></msub></mrow></mtd></mtr></mtable></mfenced> obtain The adaptive law of Among them, κ ₁ ＞0, Γ ₁ ∈R ^4×4 , λ _Ω =diag{λ ₁ ,λ ₂ ,λ ₃ ,λ ₄ }＞0 are the controller parameters, g _s , f ₁ , Ψ are the system coefficient matrix , Ψ ^T is the transpose matrix of Ψ, e ₁ is the error, The slow loop follows the derivative of the instruction, is the estimated value of the slow loop state variable, is the estimation error, δ is the sliding mode function; Step 2-4) Attitude Angle Rate Controller Design Define the error e ₂ =ω-ω _c , and derive the error <mrow><msub><mover><mi>e</mi><mo>&amp;CenterDot;</mo></mover><mn>2</mn></msub><mo>=</mo><mover><mi>&amp;omega;</mi><mo>&amp;CenterDot;</mo></mover><mo>-</mo><msub><mover><mi>&amp;omega;</mi><mo>&amp;CenterDot;</mo></mover><mi>c</mi></msub><mo>=</mo><msup><mi>&amp;Xi;</mi><mi>T</mi></msup><msub><mi>&amp;theta;</mi><mi>&amp;omega;</mi></msub><mo>+</mo><msub><mi>g</mi><mi>f</mi></msub><msub><mi>M</mi><mi>c</mi></msub><mo>+</mo><mi>D</mi><mo>-</mo><msub><mover><mi>&amp;omega;</mi><mo>&amp;CenterDot;</mi>mo></mover><mi>c</mi></msub></mrow> virtual controller As a disturbance, we get in Ξ ^T is the transposition of the system fast loop system matrix; the attitude angle rate loop control design is in is the inverse of the fast loop system state matrix g _f , Ξ ^T is the transposition of the system fast loop system matrix, θ _ω is the fast loop state variable, g _f is the system state matrix, M _c is the control torque, D is the external disturbance, is the derivative of the slow loop controller, κ ₂ is the control design parameter, and is the observed tor output, is the transpose of the slow loop state system matrix, Estimated by the observer, defining the observation error Adaptive law is Among them, κ ₂ ＞0, Γ ₂ ∈R ^17×17 ＞0, λ _ω ∈R ^17×17 ＞0 are the controller parameters, To estimate the error, the observer is designed as <mfenced open = "{" close = ""><mtable><mtr><mtd><mrow><mover><mi>d</mi><mo>^</mo></mover><mo>=</mo><mi>z</mi><mo>+</mo><mi>Q</mi><mrow><mo>(</mo><mi>e</mi><mo>)</mo></mrow></mrow></mtd></mtr><mtr><mtd><mrow><mover><mi>z</mi><mo>&amp;CenterDot;&amp;CenterDot;</mo></mover><mo>=</mo><mo>-</mo><mi>L</mi><mrow><mo>(</mo><mi>e</mi><mo>)</mo></mrow><mi>z</mi><mo>-</mo><mi>L</mi><mrow><mo>(</mo><mi>e</mi><mo>)</mo></mrow><mrow><mo>(</mo><msup><mi>&amp;Xi;</mi><mi>T</mi></msup><msub><mover><mi>&amp;theta;</mi><mo>^</mo></mover><mi>&amp;omega;</mi></mi>msub><mo>+</mo><msub><mi>g</mi><mi>f</mi></msub><mi>M</mi><mo>+</mo><mi>Q</mi><mo>(</mo><mi>e</mi><mo>)</mo></mrow><mo>)</mo></mrow></mtd></mtr></mtable></mfenced> in, q ₁ (e), q ₂ (e), . . . , q _n (e) are observer parameters. 4. the hypersonic vehicle attitude motion coordination control method based on coordination factor according to claim 1, it is characterized in that, described step 3) the concrete process following steps: With the attitude angular rate loop controller M _c and the rudder distribution matrix g _fδ , the coordination torque is expressed as M _c =g _Γfδ ·δ, where <mfenced open = "" close = ""><mtable><mtr><mtd><mrow><msub><mi>g</mi><mrow><mi>&amp;Gamma;</mi><mi>f</mi><mi>&amp;delta;</mi></mrow></msub><mo>=</mo><msub><mi>g</mi><mrow><mi>f</mi><mi>&amp;delta;</mi></mrow></msub><mo>+</mo><mi>&amp;Gamma;</mi></mrow></mtd></mtr><mtr><mtd><mrow><mo>=</mo><mfenced open = "[" close = "]"><mtable><mtr><mtd><mrow><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub><mo>+</mo><msub><mi>&amp;lambda;</mi><mi>e</mi></msub></mrow></mtd><mtd><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>p</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub></mtd></mtr><mtr><mtd><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><mi>&amp;delta;</mi><mi>e</mi></mrow></msub></mtd><mtd><mrow><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mi>mo><mi>&amp;delta;</mi><mi>a</mi></mrow></msub><mo>+</mo><msub><mi>&amp;lambda;</mi><mi>a</mi></msub></mrow></mtd><mtd><msub><mi>g</mi><mrow><mi>q</mi><mo>,</mo><mi>&amp;delta;</mi><mi>r</mi></mrow></msub></mtd></mtr><mtr><mtd><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>e</mi></msub></mrow></msub></mtd><mtd><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>a</mi></msub></mrow></msub></mtd><mtd><mrow><msub><mi>g</mi><mrow><mi>r</mi><mo>,</mo><msub><mi>&amp;delta;</mi><mi>r</mi></msub></mrow></msub><mo>+</mo><msub><mi>&amp;lambda;</mi><mi>r</mi></msub></mrow></mtd></mtr></mtable></mfenced></mrow></mtd></mtr></mtable></mfenced> in Assign matrix parameters for pitch channel rudder surface, g _q,δe , g _q,δa , g _q,δr are yaw channel rudder surface assignment matrix parameters, Assign matrix parameters to the control surface of the roll channel, λ _e , λ _a , and λ _r are coordination factors; then the control surface command can be calculated as δ=g _Γfδ ^-1 ·M _c . 5. according to claim 4, based on the hypersonic vehicle attitude motion coordination control method of coordination factor, it is characterized in that, when coordination strategy adopts aileron rudder and rudder coordination deflection to realize yaw moment and roll moment, elevator deflection realizes The pitching moment and rudder surface coordination design are: δ _coe = δ _e , δ _coa = Eδ _r , δ _cor = δ _r .