CN106842912B - Hypersonic speed maneuvering flight control surface saturation robust control method - Google Patents

Abstract

The invention discloses a hypersonic maneuvering flight control surface saturation robust control method, and belongs to flight control methods in the technical field of aerospace. The invention aims to solve the problem that the performance of a hypersonic flight vehicle (HSV) control system is poor due to amplitude saturation and compound interference of a control surface in a hypersonic maneuvering flight process, and provides a design of an anti-saturation robust control method. The method provides a novel anti-saturation auxiliary design system, the order of the system is the same as that of an attitude control system, and the system is not only suitable for a SISO system but also suitable for a MIMO system. The auxiliary system variable is introduced into a backstepping error variable, and an HSV maneuvering flight control law is designed by applying the backstepping idea, so that the closed loop stability of the system is ensured. In addition, the invention provides a hybrid tracking differentiator-based interference observer (HTDDO) for tracking and approximating the composite interference and designing a compensation control law.

Description

Robust control method for anti-rudder surface saturation in hypersonic maneuvering flight

Technical field:

The invention relates to a flight control method in the technical field of aerospace, in particular to a hypersonic maneuvering flight anti-rudder surface saturation robust control method, which is especially suitable for flight under the conditions of steering rudder surface saturation and compound interference during hypersonic maneuvering flight Control Method.

Background technique:

Hypersonic vehicles (Hypersonic Vehicles, HSV) refer to aircraft with a flight speed of Mach 5 or more, which are mainly active in the airspace from 20km to 100km above the ground. Due to the characteristics of large flight envelope, complex flight environment, high maneuverability, and multi-mission mode, HSV inevitably has uncertainties caused by its internal structure and aerodynamic parameters and interference caused by the external environment. Highly coupled, the controlled object exhibits strong nonlinear dynamic characteristics. These factors will increase the difficulty of designing the attitude control algorithm. If the classical linear control method is used, the control accuracy will be reduced or even the system will become unstable. Therefore, designing a strong and robust nonlinear control method has become a research hotspot of HSV flight control.

The maneuvering of HSV refers to the process of changing the flight state (altitude, speed, and flight direction) of HSV within a certain period of time. At present, the main research is on vertical maneuvers such as acceleration and deceleration, jumping, and diving, as well as lateral conventional maneuvers such as turning and circling. In order to complete these maneuvers, it is necessary to impose mandatory constraints on the flight state, especially the state variables such as the angle of attack, sideslip angle, and roll angular velocity; in addition, the HSV flight airspace is thin and the rudder deflection is easy to be saturated. When the actuator is saturated, the output signal of the controller is further increased, but the input of the actuator to the controlled object cannot be increased. As a result, the output of the controller is inconsistent with the actual control input of the system, which will inevitably lead to the performance of the control system. decline, or even destabilize, with serious consequences. Therefore, in order to ensure the stability of HSV maneuvering flight, anti-rudder surface saturation is the first problem to be solved in the design of maneuvering flight control method.

Due to the unique flight airspace of HSV, its maneuvering flight has model parameter uncertainty and external large torque disturbance, and the maneuvering flight control system must have strong and robust stabilization capability. Although the laboratory has obtained the aerodynamic parameters for the simulated flight of the prototype under hypersonic flight conditions through wind tunnel experiments, in fact, the near-space flight environment is complex, and various unknown factors in the real flight environment have not yet been mastered, so the aircraft There are structural and parametric uncertainties between the model and the actual aircraft. Therefore, in the process of HSV maneuvering flight, anti-jamming is also an important problem to be solved by the maneuvering flight control method.

If the uncertain disturbance items of the system can be accurately estimated, the controller can be designed to compensate for the disturbance, thereby improving the robustness of the system. Therefore, scholars at home and abroad have conducted a lot of research on disturbance observers. The disturbance observer of the tracking differentiator has better approximation and tracking ability for uncertain disturbance items. Bu X W,Wu X Y,Chen Y X,et al.Design of a class of new nonlinear disturbance observers based on tracking differentiators for uncertain dynamic systems[J]. International Journal of Control Automation & Systems, 2015, 13(3): 595-602. This paper lists various forms of current disturbance observers based on tracking differentiators, and makes a comparative analysis. Considering that the hybrid tracking differentiator is a differentiator designed based on the singular perturbation principle, it has fast convergence in the whole process, and can avoid the chattering phenomenon, so the invention proposes a hybrid differentiator-based design of a disturbance observer, which approximates the tracking ability. good. Professor JingZhou of Norwegian University of Science and Technology (Zhou, J., Wen, C.: Robust adaptive control of uncertain nonlinear systems in the presence of input saturation. In: Proceedings of 14th IFACSymposium on System Identification, Newcastle, Australia (2006)) for single input The single-output system model proposes an anti-saturation auxiliary system. The anti-saturation effect is good, but it can only be applied to the SISO system, and the model form of the control object is also different from that of the HSV model. Inspired by the anti-saturation auxiliary system, the present invention designs an anti-rudder surface saturation auxiliary control system suitable for HSV maneuvering flight.

Invention content:

The present invention focuses on the practical control problems that the control surface of the HSV is easy to be saturated and the external interference is large when the HSV maneuvers in the near space. In order to solve the problem of control surface saturation, a new anti-saturation auxiliary control system is proposed and introduced into the HSV maneuvering flight control method. This method can ensure that the closed-loop system is globally asymptotically stable, and the tracking error is clearly controllable. Aiming at the compound disturbance suffered by the aircraft, a hybrid tracking differentiator-based nonlinear disturbance observer (HTDDO) is proposed to track and approximate the disturbance, and a compensation control law is designed to suppress the disturbance. HTDDO tracking approximation works well.

The present invention adopts the following technical solutions: a robust control method for anti-rudder surface saturation for hypersonic maneuvering flight, comprising the following steps:

(1) Transform the HSV motion equation under the assumption of spherical earth into an affine nonlinear equation with limited amplitude of rudder surface for the design of maneuvering flight control law, including trajectory loop affine nonlinear equation, slow attitude Loop affine nonlinear equation and attitude fast loop affine nonlinear equation;

(2) Aiming at the affine nonlinear equation of the HSV trajectory loop constructed in step (1), according to the maneuvering flight command signal, the hypersonic flight trajectory loop control law is designed by using the backstepping method;

(3) For the compound disturbance term of the HSV attitude slow-loop and fast-loop affine nonlinear equations constructed in step (1), design a hybrid tracking differentiator-based disturbance observer (HTDDO) to approximate the compound disturbance term. ;

(4) For the two sets of affine nonlinear equations of the HSV attitude fast loop and slow loop constructed in step (1), design an anti-saturation auxiliary control system of the same order as the system;

(5) The auxiliary control system variable designed in step (2) is introduced into the error variable in the backstepping method, and the attitude control law considering the saturation of the rudder surface amplitude is deduced by applying the design idea of the backstepping method.

Further, the form of three groups of affine nonlinear equations of HSV maneuvering flight in above-mentioned step (1) is as follows:

A. Trajectory loop affine nonlinear equation

Among them, P=[χ, γ] ^T is the track control vector, χ, γ are the track azimuth and track inclination respectively; f _v =[f _χ , f _γ ] ^T is the nonlinearity of the state quantity of the track control loop function, g _v is the trajectory loop control gain matrix, the specific expression is as follows:

为动压，S为机翼有效参考面积，m为飞行器质量，V为空速，γ为航迹倾角，R为飞行器到地心的距离，χ为航迹方位角，δ为纬度，ω_E为地球旋转角速度；Here u _v = [C _Lα sinσ C _Lα cosσ] ^T is the control vector of the trajectory loop;

is the dynamic pressure, S is the effective reference area of the wing, m is the mass of the aircraft, V is the airspeed, γ is the track inclination, R is the distance from the aircraft to the center of the earth, χ is the track azimuth, δ is the latitude, ω _E is the angular velocity of the earth's rotation;

B. Attitude slow loop affine nonlinear equation

Among them, Ω = [α, β, σ] ^T attitude slow loop airflow attitude angle vector, α, β, σ are attack angle, sideslip angle and track roll angle respectively; ω _c = [p _c ,q _c ,r _c ] ^T is the angular rate tracking signal of the attitude fast loop, p, q, r are the pitch, roll and yaw angular rates respectively; d _s ∈ R ³ is the composite disturbance error of the attitude slow loop; f _s = [f _α ,f _β ,f _σ ] ^T is the nonlinear function of the attitude slow loop state vector, g _s is the attitude slow loop control gain matrix, and the specific expression is as follows:

Here, C _L,a is the basic lift coefficient, and C _C,β is the basic lateral force coefficient;

C. Attitude fast loop affine nonlinear equation

Among them, ω=[p,q,r] ^T is the angular rate vector of the attitude fast loop; δ _c =[δ _e ,δ _a ,δ _r ] ^T is the deflection angle of the aerodynamic rudder surface, δ _e ,δ _a ,δ _r respectively is the deflection angle of the left and right aileron elevators and rudders; sat(δ _c ) is the actual rudder surface deflection after the rudder amplitude is saturated, which is the final control amount of the attitude control system; d _f ∈ R ³ is the attitude fast The composite disturbance error of the loop; f _f =[f _p ,f _q ,f _r ] ^T is the nonlinear function of the attitude fast loop state vector, g _f is the attitude fast loop control gain matrix, and the specific expression is:

Here, I _x , I _y , and I _z are the moments of inertia around the x, y, and z axes, respectively; b is the wingspan; c is the average aerodynamic chord; X _cg is the distance between the center of mass and the focus,

为飞行器气动参数。

are the aerodynamic parameters of the aircraft.

Further, in the described step (3), for the composite interference term of the attitude slow loop and the fast loop, the designed HTDDO form is:

Attitude slow loop HTDDO:

Attitude fast loop HTDDO:

The specific form of the f(x ₁ , x ₂ ) function is:

sig(x) ^a = sgn(x)|x| ^a

sgn(x)=diag(sgn(x ₁ ),sgn(x ₂ ),...,sgn(x _n ))

where a ₀ , a ₁ , a ₂ , b ₀ , b ₁ , ∈ are design parameters, which are all positive values.

Further, the anti-saturation aided design system constructed in the step (2) is as follows:

where Δδ=sat(δ _c )-δ _c is the input signal of the auxiliary system; λ ₁ ∈ R ³ and λ ₂ ∈ R ³ are the state quantities of the auxiliary system, c ₁ ∈ R ^3×3 and c ₂ ∈ R ^3×3 are design parameter matrices, which are diagonal matrices whose diagonal elements are positive numbers, and c ₁ should also satisfy λ _min (c ₁ )-0.5>0.

Further, the HSV maneuvering flight control law derived by applying Lyapunov stability theory in the step (3) is as follows:

where z _v , z ₁ and z ₂ are the defined error variables,

z _v =PP _c , z ₁ =Ω-Ω _c -λ ₁ ,z ₂ =ω-ω _c -λ ₂ , k _v ∈R ^2×2 ,k ₁ ∈R ^3×3 ,k ₂ ∈R ^{3× 3} ,c ₁ ∈R ^3×3 ,c ₂ ∈R ^3×3

are design parameter matrices, which are diagonal matrices with positive diagonal elements, and c ₁ should also satisfy λ _min (c ₁ )-0.5>0, where λ _min (·) represents the minimum eigenvalue of the matrix.

The present invention has the following beneficial effects:

(1) The anti-saturation auxiliary control system proposed by the present invention has a simple structure and better performance, and gives the system instantaneous error tracking performance inequality, which provides a basis for the adjustment of the design parameters of the auxiliary system in the engineering implementation, so that it can be better improved system performance.

(2) The present invention proposes that the nonlinear disturbance observer is a hybrid tracking differentiator-based nonlinear disturbance observer (HTDDO), which has a faster approximation and tracking capability for the complex disturbances suffered by the system, and its structure is relatively simple, There are fewer design parameters. Combined with the hypersonic maneuvering flight control method of HTDDO, it has a strong adaptability to the dynamic uncertainty of the hypersonic vehicle system and the external disturbances encountered in the maneuvering flight, thereby effectively improving the robust performance of the flight control system.

Description of drawings:

Figure 1 is a block diagram of the control system structure.

Figures 2(a) and 2(b) are the running results of track azimuth and track inclination without anti-saturation auxiliary system.

Figures 3(a), 3(b) and 3(c) are the operating results of the angle of attack, sideslip angle, and roll angle without the anti-saturation auxiliary system.

Figures 4(a), 4(b) and 4(c) are the operating results of pitch, roll and yaw angular velocity without anti-saturation assist system.

Figures 5(a), 5(b) and 5(c) are the operating results of the left elevon rudder, right elevon rudder and rudder without anti-saturation assist system.

Figures 6(a) and 6(b) are the running results of the track azimuth and track inclination with the anti-saturation assistance system added.

Figures 7(a), 7(b) and 7(c) are the operating results of the angle of attack, sideslip angle and roll angle with the anti-saturation assist system added.

Figures 8(a), 8(b) and 8(c) are the operational results of the pitch, roll and yaw angular velocities with the anti-saturation assist system added.

Figures 9(a), 9(b) and 9(c) are the operating results of the left elevon rudder, right elevon rudder and rudder with the anti-saturation assist system added.

Figures 10(a), 10(b) and 10(c) are the running results of the slow-loop complex disturbance approximation tracking with the anti-saturation assist system.

Figures 11(a), 11(b) and 11(c) are the running results of the fast-loop composite disturbance approximation tracking with the anti-saturation auxiliary system added.

Detailed ways:

The following describes in detail the anti-rudder surface saturation robust control method for hypersonic maneuvering flight proposed by the present invention with reference to the embodiments and the accompanying drawings. The embodiment adopts a model of a winged-cone configuration proposed by NASA Langley Research Center as a research object.

Twelve state equations of HSV are established (Mooij E. Motion of a vehicle in a planetaryatmosphere [J]. NASA STI/Recon Technical Report N, 1994, 96: 11743). Its specific form is as follows:

The aerodynamic model of the aircraft is mainly from Keshmiri S, Colgren R, Mirmirani M.Six DoFNonlinear Equations of Motion for a Generic Hypersonic Vehicle[C].AIAAAtmospheric Flight Mechanics Conference and Exhibit,United States:AIAA,2007,1-28.

The present invention only considers the control problem of the HSV reentry maneuvering flight, and does not consider the maneuvering trajectory planning and guidance problems, so the trajectory loop state quantity does not consider the control of the speed V. The above state equations of track azimuth χ and track inclination γ are rewritten into affine nonlinear equations.

Among them, P=[χ,γ] ^T is the track control vector; u _v =[C _Lα sinσ C _Lα cosσ] ^T is the control vector of the trajectory loop; f _v =[f _χ ,f _γ ] ^T is the trajectory control loop The nonlinear function of the state quantity, g _v is the trajectory loop control gain matrix, and the expression is as follows.

According to the affine nonlinear equation of the trajectory loop, set the maneuvering flight command signal P _c =[χ _c ,γ _c ] ^T The control law designed according to the NGPC method is:

where z _v = [PP _c ] ^T , and k _v is a control parameter. u _v =[C _La sinσ,C _La cosσ] ^T , C _La is a nonlinear function related to the angle of attack α, through the Newton iteration method, the attitude control signal α _c ,σ _c required to track the maneuver command signal can be obtained , Referring to the BTT control of banked turn, to realize no-slip turn, so as to reduce the design pressure of the control surface and protection system. Set the sideslip angle control signal β _c =0, so as to obtain the complete attitude control system tracking command signal Ω _c =[α _c ,β _c ,σ _c ] ^T .

According to the state equation for establishing attitude variables, it is transformed into the attitude loop affine nonlinear equation form of HSV maneuvering flight. According to the response time of attitude variables, attitude variables can be divided into slow variables (α, β, σ) and fast variables (p ,q,r), and establish the attitude slow-loop affine nonlinear equation and attitude fast-loop affine nonlinear equation respectively. Its specific form is as follows:

A. Attitude slow loop affine nonlinear equation:

Among them, f _s =[f _α , f _β , f _σ ] ^T is the state vector nonlinear function, g _s is the control gain matrix, and the specific expression is as follows.

where Ω=[α,β,σ] ^T attitude slow loop airflow attitude angle vector, α,β,σ are the angle of attack, sideslip angle and bank angle respectively; ω _c =[p _c ,q _c , _rc ] ^T is the angular rate tracking signal of the attitude fast loop, p, q, r are the pitch, roll and yaw angular rates, respectively.

B. Attitude fast loop affine nonlinear equation:

where ω=[p,q,r] ^T is the attitude fast loop angular rate vector; δ _c =[δ _e ,δ _a ,δ _r ] ^T is the deflection angle of the aerodynamic rudder surface, δ _e ,δ _a ,δ _r are respectively The deflection angles of the left and right aileron elevators and rudders; sat(δ _c ) is the actual rudder surface deflection after the rudder surface amplitude is saturated, which is the final control amount of the attitude control system; f _f =[f _p ,f _q ,f _r ] ^T is the nonlinear function of the attitude fast loop state vector, g _f is the attitude fast loop control gain matrix, and the specific expression is as follows.

The HTDDO is designed to approach the disturbance for the compound disturbance in the attitude slow loop and fast loop.

The attitude slow loop HTDDO is designed as follows:

The attitude fast loop HTDDO is designed as follows:

The specific form of the f(x ₁ , x ₂ ) function is:

sig(x) ^a = sgn(x)|x| ^a

sgn(x)=diag(sgn(x ₁ ),sgn(x ₂ ),...,sgn(x _n ))

Here a ₀ , a ₁ , a ₂ , b ₀ , b ₁ , ∈ are design parameters, which are all positive numbers.

In order to effectively compensate for the influence of rudder surface saturation, an anti-saturation auxiliary control system is constructed, and its specific form is as follows:

where λ ₁ ∈ R ³ and λ ₂ ∈ R ³ are auxiliary system state quantities, c ₁ ∈ R ^3×3 and c ₂ ∈ R ^3×3 are design parameter matrices, which are diagonal matrices with positive diagonal elements , and c ₁ should also satisfy λ _min (c ₁ )-0.5>0, where λ _min (·) represents the minimum eigenvalue of the matrix. Δδ=sat(δ _c )-δ _c is the input signal of the auxiliary system.

Define the error variables z ₁ =Ω-Ω _c -λ ₁ and z ₂ =ω-ω _c -λ ₂ . If the rudder surface does not saturate, that is, Δδ=0, and the auxiliary system state quantities λ ₁ and λ ₂ are zero, the auxiliary system does not affect the error vector.

(1) Taking the derivative of z ₁ , we get:

求导得：(2) Consider the Lyapunov function

Obtain:

(3) Design the slow loop control law:

中得：Substitute into

Win:

The first term is negative definite, and the second term can be eliminated in the next step.

(4) Taking the derivative of z ₂ , we get:

(5) Consider the amplified Lyapunov function V=V ₁ +0.5z ₂ ^T z ₂ to derive:

Design the slow loop control law:

中得：Substitute into

Win:

Finally, the control laws of the fast loop and the slow loop are obtained as follows:

According to the derivation method of the above control law, the state tracking error can satisfy:

Simultaneous attitude loop affine nonlinear equation and auxiliary design system equation can deduce that the transient tracking error performance of airflow attitude angle satisfies:

可以根据以上不等式调节控制器参数，以改善控制器性能。Where ||·|| ₂ represents the L ₂ norm of the state quantity, which is specifically defined as:

The controller parameters can be adjusted according to the above inequalities to improve the controller performance.

The present invention is simulated and verified in the MATLAB2014a environment, and the initial flight status is as follows: the height H=35km, the flight speed V=3000m/s, the aircraft mass is 136820kg, and the rudder surface limit is ±28°. The initial attitude angle and angular rate are: α ₀ =3.0°, β ₀ =0°, σ ₀ =2.5°, p ₀ =q ₀ =r ₀ =0rad/s. The maneuvering flight command signal P _c =[χ _c ,γ _c ] ^T is shown in Figs. 5(a) and 5(b). The simulation design parameters are shown in the table below. The attitude slow loop composite disturbance is d _s =[d _s1 ,d _s2 ,d _s3 ] ^T , where d _s1 =0.001sin(t+1)cos(2t), _ds2 =0.003·cos(t+1)sin( 2t+2), d _s3 =0.005sin(t+1)sin(2t), the attitude fast loop composite interference term d _f =[d _f1 ,d _f2 ,d _f3 ] ^T , where d _f1 =0.05sin(t+ 1), d _f2 =0.04cos(2t+2), d _f3 =0.03sin(t+1).

In order to highlight the role of the auxiliary system in the HSV maneuvering flight, the present invention provides two sets of simulation results. Figures 2(a) to 5(c) are the results of the maneuvering flight without the anti-saturation auxiliary control system, and Figure 6 (a) to Figure 11 (c) are the results of the operation of the anti-saturation auxiliary control system.

The above are only the preferred embodiments of the present invention. It should be pointed out that for those skilled in the art, several improvements can be made without departing from the principles of the present invention, and these improvements should also be regarded as the invention. protected range.

Claims (4)

1. a hypersonic maneuvering flight anti-rudder surface saturation robust control method, is characterized in that: comprise the following steps (1) Transform the HSV motion equation under the assumption of spherical earth into an affine nonlinear equation with limited amplitude of rudder surface for the design of maneuvering flight control law, including trajectory loop affine nonlinear equation, slow attitude Loop affine nonlinear equation and attitude fast loop affine nonlinear equation; (2) Aiming at the affine nonlinear equation of the HSV trajectory loop constructed in step (1), according to the maneuvering flight command signal, the hypersonic flight trajectory loop control law is designed by using the backstepping method; (3) For the compound disturbance term of the HSV attitude slow-loop and fast-loop affine nonlinear equations constructed in step (1), design a disturbance observer based on a hybrid tracking differentiator to approximate the compound disturbance term; (4) For the two sets of affine nonlinear equations of the HSV attitude fast loop and slow loop constructed in step (1), design an anti-saturation auxiliary control system of the same order as the system; (5) Introduce the auxiliary control system variable designed in step (4) into the error variable in the backstepping method, apply the design idea of the backstepping method, and design an attitude control law considering the saturation of the rudder surface amplitude; The forms of the three groups of affine nonlinear equations for HSV maneuvering flight in the step (1) are as follows: A. Trajectory loop affine nonlinear equation

Among them, P=[χ, γ] ^T is the track control vector, χ, γ are the track azimuth and track inclination respectively; f _v =[f _χ , f _γ ] ^T is the nonlinearity of the state quantity of the track control loop function, g _v is the trajectory loop control gain matrix, the specific expression is as follows:

Among them, u _v =[C _Lα sinσ C _Lα cosσ] ^T is the control vector of the trajectory loop, and C _La is a nonlinear function of the angle of attack α;

is the dynamic pressure, S is the effective reference area of the wing, m is the mass of the aircraft, V is the airspeed, γ is the track inclination, R is the distance from the aircraft to the center of the earth, χ is the track azimuth, δ is the latitude, ω _E is the angular velocity of the earth's rotation; B. Attitude slow loop affine nonlinear equation

Among them, Ω=[α,β,σ] ^T is the attitude angle vector of the air flow in the attitude slow loop, α, β, σ are the angle of attack, sideslip angle and track roll angle respectively; ω _c =[p _c ,q _c , rc ] ^T is the angular rate tracking signal of the attitude fast loop, p, q, _r are the roll, pitch and yaw angular rates, respectively, p _c , q _c , rc are the roll, pitch and yaw rate of the attitude fast loop, respectively _. The tracking signal of the yaw rate; d _s ∈ R ³ is the composite disturbance error of the attitude slow loop; f _s = [f _α , f _β , f _σ ] ^T is the nonlinear function of the state vector of the attitude slow loop, g _s is the attitude The slow loop control gain matrix, the specific expression is as follows:

Among them, C _L,a is the basic lift coefficient, and C _C,β is the basic lateral force coefficient; C. Attitude fast loop affine nonlinear equation

Among them, ω=[p,q,r] ^T is the angular rate vector of the attitude fast loop; δ _c =[δ _e ,δ _a ,δ _r ] ^T is the deflection angle of the aerodynamic rudder surface, δ _e ,δ _a ,δ _r respectively is the deflection angle of the left and right aileron elevators and rudders; sat(δ _c ) is the actual rudder surface deflection after the rudder amplitude is saturated, which is the final control amount of the attitude control system; d _f ∈ R ³ is the attitude fast The composite disturbance error of the loop; f _f =[f _p ,f _q ,f _r ] ^T is the nonlinear function of the attitude fast loop state vector, g _f is the attitude fast loop control gain matrix, and the specific expression is:

Among them, I _x , I _y , and I _z are the moments of inertia around the x, y, and z axes of the body, respectively; b is the wingspan; c is the average aerodynamic chord length; X _cg is the distance between the center of mass and the focus,

is the aerodynamic parameter of the aircraft;

are the aerodynamic parameters of the yaw moment caused by the left aileron elevator; C _l,β , C _m,α , C _n,β represent the aerodynamic torque parameters caused by the airflow attitude angle of the slow attitude loop; C _D,α represent the slow attitude loop The aerodynamic drag parameters caused by the angle of attack; C _l,p , C _l,r , C _m,q , C _n,p , C _n,r represent the aerodynamic torque parameters caused by the angular rate of the attitude fast loop. 2. hypersonic maneuvering flight anti-rudder surface saturation robust control method according to claim 1, is characterized in that: in described step (3), for the composite interference term of attitude slow loop and fast loop, the HTDDO form of design is : Attitude slow loop HTDDO:

in,

is the composite disturbance error of the attitude slow loop HTDDO; Attitude fast loop HTDDO:

The specific form of the f(x ₁ , x ₂ ) function is:

sig(x) ^a = sgn(x)|x| ^a sgn(x)=diag(sgn(x ₁ ),sgn(x ₂ ),...,sgn(x _n )) where a ₀ , a ₁ , a ₂ , b ₀ , b ₁ , ∈ are design parameters, which are all positive values. 3. hypersonic maneuvering flight anti-rudder surface saturation robust control method according to claim 1, is characterized in that: the anti-saturation auxiliary design system constructed in described step (2) is as follows:

where Δδ=sat(δ _c )-δ _c is the input signal of the auxiliary system; λ ₁ ∈ R ³ and λ ₂ ∈ R ³ are the auxiliary state quantities, c ₁ ∈ R ^3×3 and c ₂ ∈ R ^3×3 are The diagonal elements of the design are diagonal matrices of positive numbers, and c ₁ should also satisfy λ _min (c ₁ )-0.5>0, where λ _min (·) represents the minimum eigenvalue of the matrix. 4. hypersonic maneuvering flight anti-rudder surface saturation robust control method according to claim 3, is characterized in that: in described step (3), the HSV maneuvering flight control law that applies Lyapunov stability theory deriving is as follows:

Among them, z _v , z ₁ and z ₂ are the defined error variables, z _v =PP _c , z ₁ =Ω-Ω _c -λ ₁ , z ₂ =ω-ω _c -λ ₂ ; k _v ∈R ^{2× 2} , k ₁ ∈ R ^3×3 , k ₂ ∈ R ^3×3 , c ₁ ∈ R ^3×3 , c ₂ ∈ R ^3×3 are design parameter matrices, which are diagonal matrices with positive diagonal elements , and c ₁ should satisfy λ _min (c ₁ )-0.5>0, where λ _min (·) represents the minimum eigenvalue of the matrix.

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