Aerodynamic servo elastic active control system and method for air-breathing hypersonic aircraft

Abstract

The invention provides a pneumatic servo elastic active control system and method for an air-breathing hypersonic aircraft, and the method comprises the steps: an elastic frequency estimator obtains an elastic frequency estimation value of a controlled object based on the pitch angle information transmitted by the controlled object, and transmits the elastic frequency estimation value to a state disturbance joint estimator; the state disturbance joint estimator obtains a corresponding state estimation result, a disturbance estimation result and a pitch angle estimation result based on the received pitch angle information, elevator deflection angle and elastic frequency estimation value; and the attitude controller obtains the elevator deflection angle of the controlled object based on the received rigid body pitch angle control value, the state estimation result and the disturbance estimation result, and sends the elevator deflection angle to the state-disturbance joint estimator and the controlled object. According to the invention, the controlled object can be stably controlled.

Description

Technical Field

[0001] The present invention relates to the field of air-breathing hypersonic aircraft control, and in particular to an aerodynamic servoelastic active control system and method for an air-breathing hypersonic aircraft. Background Art

[0002] At present, most of the studies on the control of elastic air-breathing hypersonic vehicles assume that the rigid body motion information can be measured and only consider the aeroelastic effect. However, the sensor measurement signal contains both rigid body motion information and elastic vibration information, and the rigid body motion information cannot be directly obtained. Therefore, in addition to aeroelasticity, the controller design must also consider the servoelasticity caused by the elastic vibration feedback in the measurement signal, and the two simultaneously form a complex aeroservoelastic effect. The aeroservoelastic control methods of elastic air-breathing hypersonic vehicles mainly include robust control, amplitude stability, phase stability and active control. Among them, robust control characterizes the desired elastic mode suppression performance through the closed-loop performance transfer function, and obtains the controller by solving the optimization problem. The idea of ​​amplitude stability is to passively suppress the elastic mode by designing a notch filter. Compared with amplitude stability, phase stability can expand the rigid body control bandwidth and obtain a satisfactory stability margin. Phase stability optimizes the position of the sensor and the actuator to realize the self-stabilization of the elastic mode by directly feeding the measurement signal into the flight control system. It belongs to the semi-active control method with fixed elastic state feedback gain. In order to quickly suppress multiple elastic modes simultaneously, it is necessary to study active control methods that can individually design the feedback gain of each elastic state. Summary of the invention

[0003] In view of the above technical problems, the technical solution adopted by the present invention is:

[0004] According to a first aspect of the present invention, an aerodynamic servoelastic active control system for an air-breathing hypersonic aircraft is provided, the system comprising: an elastic frequency estimator, a state-disturbance joint estimator and an attitude controller, wherein the elastic frequency estimator is used to obtain an elastic frequency estimation value of the controlled object based on pitch angle information sent by the controlled object by using an elastic frequency estimation model, and send the elastic frequency estimation value to the state-disturbance joint estimator; the state-disturbance joint estimator is used to estimate a rigid body state, an elastic body state, a disturbance factor of the rigid body state, a disturbance factor of the elastic body state and pitch angle information by using a state-disturbance joint estimation model based on received pitch angle information, elevator deflection angle and elastic frequency estimation value, obtain corresponding state estimation results, disturbance estimation results and pitch angle estimation results, and send the state estimation results and disturbance estimation results to the attitude controller; the attitude controller is used to obtain the elevator deflection angle of the controlled object by using an attitude control model based on the received rigid body pitch angle control value, state estimation results and disturbance estimation results, and send the elevator deflection angle to the state-disturbance joint estimator and the controlled object.

[0005] According to a second aspect of the present invention, there is provided a method for active aero-servoelastic control of an air-breathing hypersonic vehicle, comprising the following steps:

[0006] S100, in response to receiving the rigid body pitch angle control value at the current moment, the attitude controller obtains the elevator deflection angle of the controlled object by using the attitude control model, and sends it to the state disturbance joint estimator and the controlled object.

[0007] S200, the controlled object performs a corresponding operation based on the received elevator deflection angle, and obtains pitch angle information after performing the corresponding operation, and sends the information to an elastic frequency estimator.

[0008] S300: Based on the received pitch angle information, the elastic frequency estimator obtains an elastic frequency estimation value of the controlled object using an elastic frequency estimation model, and sends the elastic frequency estimation value to the state disturbance joint estimator.

[0009] S400, based on the received pitch angle information, elevator deflection angle and elastic frequency estimation value, estimate the rigid body state, elastic body state, the disturbance factor of the rigid body state, the disturbance factor of the elastic body state and the pitch angle information using the state-disturbance joint estimation model to obtain corresponding state estimation results, disturbance estimation results and pitch angle estimation results, and send the state estimation results and disturbance estimation results to the attitude controller; execute S100.

[0010] The present invention has at least the following beneficial effects:

[0011] The active control system and method for aerodynamic servoelasticity of an air-breathing hypersonic aircraft provided by the embodiment of the present invention are not strongly dependent on the model accuracy compared with robust control. Specifically, disturbance factors are introduced into the nominal observation model, and these disturbance factors are model errors, including factors such as parameter uncertainty, unmodeled dynamics, and external interference that are not considered in the process of establishing the nominal observation model. The present invention first estimates the disturbance factor through a state disturbance joint estimator, and then compensates for the disturbance factor in the attitude controller to eliminate the influence of the disturbance factor. Therefore, even when the model has strong uncertainty, the designed control system can still make the system stable. In addition, the elastic frequency can be estimated in real time, which reduces the conservatism of the controller. Compared with amplitude stability, the present invention adopts an active controller without using a notch filter, avoiding the lag phase introduced by the notch filter, so as not to affect the dynamic performance and stability margin of the rigid body control. Compared with phase stability, the present invention can suppress multiple elastic modes at the same time through the control parameters in the attitude controller, and can make the elastic mode achieve the desired suppression speed by flexibly designing the closed-loop damping of each elastic mode.

[0012] It should be understood that the contents described in this section are not intended to identify the key or important features of the embodiments of the present invention, nor are they intended to limit the scope of the present invention. Other features of the present invention will become easily understood through the following description. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for use in the description of the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0014] Figure 1 A structural block diagram of an aerodynamic servoelastic active control system for an air-breathing hypersonic aircraft provided in an embodiment of the present invention;

[0015] Figure 2 A flow chart of an active aero-servoelastic control method for an air-breathing hypersonic vehicle provided in an embodiment of the present invention;

[0016] Figure 3 is a schematic diagram of a frequency estimation result of a frequency estimator;

[0017] Figures 4a to 4e It is a schematic diagram of the estimation results of the state disturbance joint estimator;

[0018] Figure 5 Schematic diagram of the control quantities calculated for the attitude controller;

[0019] Figure 6 It is a schematic diagram of the measurement values ​​of each sensor obtained after the control quantity is input into the proxy model;

[0020] Figure 7 Schematic diagram of the tracking error of the pitch angle of a rigid body;

[0021] Figures 8a to 8c Schematic diagram of the generalized coordinate derivatives of the first three elastic modes. DETAILED DESCRIPTION

[0022] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0023] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as those commonly understood by those skilled in the art of the present invention. The terms used herein in the specification of the present invention are only for the purpose of describing specific embodiments and are not intended to limit the present invention. The term "and / or" used herein includes any and all combinations of one or more related listed items.

[0024] It should be noted that some exemplary embodiments are described as processes or methods depicted as flow charts. Although the flow charts describe the steps as sequential processes, many of the steps can be implemented in parallel, concurrently or simultaneously. In addition, the order of the steps can be rearranged. The process can be terminated when its operation is completed, but it can also have additional steps not included in the accompanying drawings. The process can correspond to a method, function, procedure, subroutine, subprogram, etc.

[0025] This embodiment provides an active aerodynamic servoelastic control system for an air-breathing hypersonic aircraft, such as Figure 1 As shown, the system includes: an elastic frequency estimator 1, a state disturbance joint estimator 2 and a posture controller 3.

[0026] The elastic frequency estimator 1 is used to obtain the elastic frequency estimation value of the controlled object based on the pitch angle information sent by the controlled object by using the elastic frequency estimation model, and send it to the state disturbance joint estimator 2. The state disturbance joint estimator 2 is used to estimate the rigid body state, elastic body state, the disturbance factor of the rigid body state, the disturbance factor of the elastic body state and the pitch angle information by using the state disturbance joint estimation model based on the received pitch angle information, elevator deflection angle and elastic frequency estimation value, obtain the corresponding state estimation result, disturbance estimation result and pitch angle estimation result, and send the state estimation result and disturbance estimation result to the attitude controller. The attitude controller 3 is used to obtain the elevator deflection angle of the controlled object by using the attitude control model based on the received rigid body pitch angle control value, state estimation result and disturbance estimation result, and send it to the state disturbance joint estimator and the controlled object.

[0027] In an embodiment of the present invention, the controlled object may be a real flying air-breathing hypersonic aircraft or a proxy model of an air-breathing hypersonic aircraft, that is, the active control system provided in an embodiment of the present invention can not only control the attitude of the actual flying aircraft but also be used for attitude control simulation analysis of the aircraft.

[0028] Furthermore, in an embodiment of the present invention, the proxy model of the elastic air-breathing hypersonic aircraft is modeled as a free beam, so that the rigid body mode and the elastic mode satisfy the orthogonality, that is, the rigid body mode and the elastic mode only affect each other through the action of force and moment. In order to improve the non-minimum phase characteristics of the ballistic inclination angle, the proxy model also introduces the canard rudder deflection angle. Specifically, the proxy model of the air-breathing hypersonic aircraft meets the following conditions:

[0029]

[0030] δ c = rδ e (3);

[0031]

[0032]

[0033] Where V represents the speed of the aircraft, is the first-order derivative of V, γ represents the ballistic inclination angle of the aircraft, is the first-order derivative of γ, h represents the height, is the first-order derivative of h, α represents the angle of attack of the aircraft, is the first-order derivative of α, Q represents the pitch angular velocity of the aircraft, is the first-order derivative of Q, Indicates the pitch angle of the aircraft. for The first derivative of i represents the generalized coordinates of the i-th elastic mode among the first three elastic modes, where i ranges from 1 to 3. is η i The first derivative of is η i The second derivative of ξ i represents the damping of the i-th elastic mode, ω i represents the true elastic frequency of the i-th elastic mode. G represents the mass of the aircraft; I yy represents the moment of inertia of the aircraft; T represents the thrust of the aircraft, D represents the drag of the aircraft, L represents the lift of the aircraft, M represents the pitch moment of the aircraft, and N represents the i represents the generalized force of the i-th order elastic mode, q represents the flight pressure of the aircraft, S represents the reference area of ​​the aircraft, represents the longitudinal reference length of the aircraft, δ e is the elevator deflection angle, δ c is the canard rudder deflection angle, φ represents the aircraft's fuel equivalence ratio, C T η represents the coupling coefficient vector from the thrust T to the generalized coordinate η, zT represents the coupling coefficient from thrust T to pitch moment M, C T，φ (·), C L (·), C D (·), C M (·), C T (·) are aerodynamic parameters, N i b Indicates b to N i The influence coefficient of b belongs to (α 2 , α, δ e , δ c ,η),N i 0 N i The corresponding constant, Denotes δ e The influence coefficient on L, Denotes δ c The influence coefficient on L; Indicates the pitch angle measured by the pitch angle sensor, represents the elastic pitch angle, represents the vibration mode of the i-th elastic mode φ i The first derivative of (p), where p represents the distance from the pitch angle sensor to the head of the aircraft.

[0034] It should be noted that, in the embodiment of the present invention, although the design process is introduced by taking the first three-order elastic modes as an example, it can be extended to the suppression of any number of elastic modes.

[0035] The expression of aerodynamic coefficient is as follows:

[0036]

[0037] Among them, the expression in the expression Represents the influence coefficient of * on ·.

[0038] It can be seen from the above statements (2) and (6) that thrust, drag, lift and pitching moment are all affected by the elastic mode, and the generalized force that excites the elastic mode is affected by the rigid body attack angle. These effects together produce the coupling problem between aerodynamics, thrust and elasticity. In order to eliminate the influence of the elevator deflection angle on lift, the present invention converts the canard rudder deflection angle and the elevator deflection angle in proportion, that is, converts them according to the above formula (3).

[0039] In the embodiment of the present invention, only attitude control, altitude and speed open-loop control of the aircraft are considered, and for this purpose, the fuel equivalence ratio φ is set to a maximum value, such as 1.5.

[0040] In the embodiment of the present invention, the pitch angle sensor may be a rate gyro.

[0041] Furthermore, the elastic frequency estimation model satisfies the following conditions:

[0042]

[0043] θ(t)=θ(t-ΔT)+(1-σ)R -1 (t)ψ θ (t)e(t);

[0044] R(t)=σR(t-ΔT)+(1-σ)[ψ θ (t)ψ θ (t) T +βI];

[0045] Among them, v(t) is the input of the elastic frequency estimation model at time t, e(t) is the output of the elastic frequency estimation model at time t, and are the pitch angles measured by the d-th pitch angle sensor and the k-th pitch angle sensor, respectively. The values ​​of k and d are 1 to m, and k≠d. m is the number of pitch angle sensors. represents the vibration mode φ of the i-th order elastic mode corresponding to the k-th pitch angle sensor i (p k ), represents the vibration mode φ of the i-th elastic mode corresponding to the d-th pitch angle sensor i (p d ), the kth pitch angle sensor and the dth pitch angle sensor can be two sensors arbitrarily selected from the m pitch angle sensors.

[0046] Among them, θ i (t) is The estimation function of is the estimated value of the elastic frequency of the i-th elastic mode among the first three elastic modes used at time t, i ranges from 1 to 3, ε is the preset adjustment parameter, z -1 and z -2 is the delay operator, β is the step adjustment factor; θ(t) is the estimation function of the elastic frequency at time t, is the estimated value of the elastic frequency of the first three elastic modes used at time t, △T is the sampling time; σ is the forgetting factor; R(t) is the covariance matrix at time t, ψ θ (t) is the gradient of e(t) with respect to θ(t), and I is the identity matrix;

[0047] In the embodiment of the present invention, β is used to affect the stability and convergence speed of parameter update and can be a preset fixed value. The larger ε is, the slower the convergence speed of the frequency estimator is and the higher the estimation accuracy is. -1 and z -2 Denotes delaying the signal by 1 sampling time and 2 sampling times respectively. i The initial value of (t) can be 0. θ (t) can be obtained based on existing methods.

[0048] In practical application scenarios, the elastic frequency ω i There are great changes in the flight process, and if it is taken as a fixed value, adaptive suppression of elastic vibration cannot be achieved. For this reason, the present invention uses a frequency estimator to estimate the elastic frequency in real time, thereby enabling adaptive suppression of elastic vibration.

[0049] Furthermore, the state disturbance joint estimation model can be obtained through the following steps:

[0050] Step 1: Construct a nominal observation model.

[0051] In the embodiment of the present invention, the nominal observation model is the basis for selecting the number and location of pitch angle sensors, and is also the basis for designing the state disturbance joint estimator. In the embodiment of the present invention, the following linear small disturbance model of attitude angle is constructed based on the aforementioned proxy model:

[0052]

[0053] Among them, x s is the state of the linear small perturbation model of the attitude angle, u2 is the total disturbance of the linear small disturbance model of the attitude angle, y is the pitch angle information of the linear small perturbation model strategy of the attitude angle, is the pitch angle measured by the hth pitch angle sensor. Q is the perturbation factor of the derivative of Q, for The disturbance factor of the derivative of , for this reason, the control system in the embodiment of the present invention allows the object model to have strong uncertainty, and its influence can be eliminated by estimating the total disturbance. A0 is the state matrix of the linear small disturbance model of the attitude angle, B1 and B2 are the input state matrices of the linear small disturbance model of the attitude angle, and C0 is the output matrix of the linear small disturbance model of the attitude angle, where:

[0054]

[0055] Among them, the expression of each element in each matrix is:

[0056]

[0057] in, Represents η i The generalized force N for the i-th elastic mode i The influence coefficient of i is the damping of the i-th elastic mode, p k represents the distance between the kth pitch angle sensor and the head of the aircraft, represents the vibration mode φ of the i-th order elastic mode corresponding to the k-th pitch angle sensor i (p k ) is the first-order derivative of .

[0058] The complete attitude angle small perturbation model needs to include the dynamics of α, but it is proved that α is unobservable. Therefore, the attitude angle small perturbation model does not include the dynamics of α, and incorporates the influence of α on Q into d Q middle.

[0059] Since the total disturbance vector u2 is unknown, in order to estimate u2 at the same time, it is expanded into the system state. The expanded system state is The expanded system model is reformulated as

[0060]

[0061] Where A is the state matrix, C is the output matrix, C=[C0 0 m×4 ]. The expanded system model is the nominal observation model.

[0062] Step 2: construct a state disturbance joint estimation model based on the aforementioned nominal observation model and elastic frequency estimation model.

[0063] In the embodiment of the present invention, the constructed state disturbance joint estimation model satisfies the following conditions:

[0064]

[0065] in, is the state of the model jointly estimated for the state disturbance, for The derivative of is the rigid body pitch angle The estimated value of is the estimated value of the rigid body pitch angular velocity Q, is the generalized coordinate η of the i-th elastic mode i The estimated value of is the generalized coordinate η of the first i-order elastic mode i The derivative of The estimated value of is the perturbation factor d of the derivative of Q Q The estimated value of for The perturbation factor of the derivative of The estimated value of is the pitch angle estimation result, is the pitch angle measured by the hth pitch angle sensor The estimated value of h ranges from 1 to m; δ e is the elevator deflection angle, y is the pitch angle information, E is the gain of the state disturbance joint estimation model, which can be configured The extreme points, i.e. the eigenvalues, are determined. B1 and C are the state matrix, input matrix and output matrix respectively; where,

[0066] In the embodiment of the present invention, the number of pitch angle sensors and their position configuration need to ensure that the state disturbance joint estimation model is observable, that is, it can be used to determine x through y. a To analyze observability, we first need to determine After calculation, It has 6 eigenvalues ​​of 0 and 6 non-zero eigenvalues, which are

[0067] Furthermore, the necessary and sufficient conditions for the state disturbance joint estimation model to be observable are:

[0068]

[0069] Among them, λ j for The eigenvalue of , j ranges from 1 to 12, rank() represents the rank of the matrix, 0 m×4 represents an all-zero matrix with m rows and m columns. That is, m needs to satisfy the above conditions.

[0070] Below, two situations are discussed:

[0071] Case 1: When λ j = 0, we can get

[0072]

[0073] Where, M1=[m1,m2,.......,m 12 ]. In order to calculate the rank of M1, a series of elementary column transformations are performed on M1. First, the order of the first 8 columns of M1 is adjusted. The adjusted M1 is:

[0074] M1=[m1 m3 m5 m7 m2 m4 m6 m8 m9 … m 12 ], then A0 and C0 become: C0=[C1 0 m×4 ],

[0075] Then, use the last 4 columns of M1 to convert the elements of the 2nd, 4th, 6th, and 8th rows of A0 to 0, that is, A0 becomes A0=[0 8×4 A1], where At this point, M1 has been transformed into At this time, rank(M)=rank(C1)+rank([A1 B2])=rank(C1)+8. In order to meet the above conditions, C1 needs to be full rank. Therefore, m≥4 is required, that is, the minimum number of pitch angle sensors is 4. In addition, the position configuration of the pitch angle sensors needs to meet the full rank requirement.

[0076] Case 2: When λ j ≠0, the above condition is equivalent to rank(N)=8, Let N = [n1 n2…n8]. In order to calculate the rank of N, we perform a series of elementary row and column transformations on N. First, adjust the order of the first 8 columns of N. After adjustment, N is: N = [n1 n3 n5 n7 n2 n4 n6 n8]. At this time, A0-λ j I 8×8 and C0 becomes:

[0077] C0=[C10 m×4 ]. Since C1 is column full rank, it can be transformed into

[0078] Then, after a series of elementary row transformations, we get Obviously, rank(N)=8, which meets the observability requirement.

[0079] In summary, the full rank of C1 is a necessary and sufficient condition to ensure that the joint estimation model of state disturbances is observable.

[0080] In the embodiment of the present invention, the positions of each sensor are further determined based on maximizing the observability matrix of the key elastic modal. Let p = [p1 p2Λp m ] represents the position vector of m (m ≥ 4) sensors, and calculates the matrix corresponding to the jth eigenvalue The condition number cond(O j ), the higher the condition number, the easier it is for the matrix to become non-full rank and the easier it is for the observability condition to be violated. Therefore, the choice of p can be solved by solving the following optimization problem:

[0081]

[0082] J(p) is the position objective function, and P represents the feasible interval of the pitch angle sensor position, that is, the position where the pitch angle sensor can actually be installed in the aircraft.

[0083] Since the maximization and minimization problems of the condition number are non-convex, numerical optimization algorithms are generally used to solve p. The present invention uses a genetic algorithm to solve p. First, a population is initialized, and each individual in the population corresponds to a candidate solution for p. Then the fitness of each individual is evaluated, and individuals with high fitness are selected for crossover and mutation. Repeat the iteration until the stop condition is met. At this point, the number and location of the sensors have been determined, and the output matrix C in the state disturbance joint estimation model can be obtained.

[0084] Furthermore, the posture control model satisfies the following conditions:

[0085]

[0086] Among them, k p , k d and To preset control parameters, It is the rigid body pitch angle control value.

[0087] For the convenience of analysis, it is assumed that there is a frequency band separation between the rigid body dynamics and the elastic body dynamics of the air-breathing hypersonic vehicle, δ e,1 and δ e,2 They are used to control rigid bodies and elastic bodies respectively. e,1 Bring in The kinetic equation can be obtained:

[0088]

[0089] Assuming that the estimation errors of the state and disturbance are close to 0, k p >0,k d >0, then will converge to 0 asymptotically. Similarly, e,2 Substitute into η i The dynamic equation of ignoring the coupling between elastic modes of different orders and the estimated error of the state can be obtained:

[0090]

[0091] It can be seen that by introducing the feedback of the elastic modal generalized coordinate derivative, the damping of the elastic vibration is increased, which can accelerate the rapid decay of the elastic vibration until Tends to 0.

[0092] In the embodiment of the present invention, the parameters in the controller are The suppression speed of the i-th order elastic mode can be changed, and different elastic modes can be designed separately according to task requirements.

[0093] Based on the same inventive concept, an embodiment of the present invention provides an active aerodynamic servoelastic control method for an air-breathing hypersonic vehicle, such as Figure 2 As shown, the following steps may be included:

[0094] S100, in response to receiving the rigid body pitch angle control value at the current moment, the attitude controller obtains the elevator deflection angle of the controlled object by using the attitude control model, and sends it to the state disturbance joint estimator and the controlled object.

[0095] In an embodiment of the present invention, the controlled object may be a real flying air-breathing hypersonic aircraft or a proxy model of an air-breathing hypersonic aircraft, that is, the active control system provided in an embodiment of the present invention can not only control the attitude of the actual flying aircraft but also be used for attitude control simulation analysis of the aircraft.

[0096] S200, the controlled object performs a corresponding operation based on the received elevator deflection angle, and obtains pitch angle information after performing the corresponding operation, and sends the information to an elastic frequency estimator.

[0097] In the embodiment of the present invention, when the controlled object is an aircraft, since the canard rudder deflection angle can be converted based on the elevator deflection angle, the elevator deflection angle of the aircraft can be controlled by the elevator deflection angle control mechanism of the aircraft to be the received elevator deflection angle, and the canard rudder deflection angle of the aircraft can be controlled by the canard rudder deflection angle control mechanism of the aircraft to be the received canard rudder deflection angle. Afterwards, the pitch angle sensor on the aircraft will measure the pitch angle at this time and send it to the elastic frequency estimator. When the controlled object is a proxy model of the aircraft, the canard rudder deflection angle can be calculated based on the received elevator deflection angle, and then the elevator deflection angle, the canard rudder deflection angle and the fuel equivalence ratio are used as inputs of the proxy model, and the corresponding pitch angle information and each state in the model can be obtained through the proxy model. The initial value of each state is determined based on the initial state of the controlled object, for example, the initial flight altitude can be 10km, etc.

[0098] S300: Based on the received pitch angle information, the elastic frequency estimator obtains an elastic frequency estimation value of the controlled object using an elastic frequency estimation model, and sends the elastic frequency estimation value to the state disturbance joint estimator.

[0099] S400, based on the received pitch angle information, elevator deflection angle and elastic frequency estimation value, estimate the rigid body state, elastic body state, the disturbance factor of the rigid body state, the disturbance factor of the elastic body state and the pitch angle information using the state-disturbance joint estimation model to obtain corresponding state estimation results, disturbance estimation results and pitch angle estimation results, and send the state estimation results and disturbance estimation results to the attitude controller; execute S100.

[0100] The control method of this embodiment can be implemented by the aforementioned control system. Therefore, the specific implementation of the method can refer to the embodiment shown in the aforementioned system and will not be described in detail.

[0101] (Example)

[0102] (1) Some parameters in the control association elastic proxy model of a certain air-breathing hypersonic vehicle are as follows. Based on the following parameters, all parameters in the state disturbance joint estimation model can be calculated:

[0103] ω1=21.17, ω2=53.92, ω3=109.1, ξ1=ξ2=ξ3=0.005,

[0104] I yy =86722.54, q=2000, S=17, Φ=1.5,r=-0.7964,

[0105]

[0106] z T =8.36,

[0107]

[0108] The present invention uses four pitch angle sensors, and the arrangement of the positions of the sensors meets the observable condition. The vibration mode derivatives of the elastic mode at the positions of the sensors are as follows:

[0109]

[0110] (2) Based on the above parameters, the parameters of the state disturbance joint estimation model can be calculated as follows:

[0111]

[0112] a 11 =0.0068, a 12 =0.0078, a 13 =0.0030,

[0113] a 21 =0.0048, a22 =-0.0010, a 23 =-0.0116,

[0114] a 31 =0.0030, a 32 =-0.0057, a 33 =0.01153,

[0115] a 41 =0.0013, a 42 =-0.0037, a 43 =0.0130

[0116] (3) Comprehensively considering the impact of various design parameters in the frequency estimator on the convergence speed and estimation accuracy, the design parameters are selected as follows:

[0117] ΔT=0.0001, ε=0.02, σ=0.7, β=1.0×10 -4 .

[0118] By configuration The eigenvalue of is -30, and the design parameter E of the state disturbance joint estimator can be calculated as:

[0119]

[0120] The design parameters in the control law are set as follows:

[0121] k p =30, k d =15,

[0122] According to the above configuration parameters, the frequency estimation result of the frequency estimator can be Figure 3 As shown, according to Figure 3 It can be seen that the frequency estimate can quickly converge to the true value. The estimation result of the state disturbance joint estimator is as follows: Figures 4a to 4e As shown, according to Figures 4a to 4e It can be seen that the error between the estimated result of the state disturbance joint estimator and the true result is very small and almost equal. The control quantity calculated by the attitude controller can be expressed as Figure 5 As shown in the figure, the measured values ​​of each sensor obtained after the control quantity is input into the proxy model can be expressed as Figure 6 As shown in the figure, the tracking error of the rigid body pitch angle is Figure 7 As shown, the generalized coordinate derivatives of the first three elastic modes are as follows: Figures 8a to 8c As shown. Figure 5 It can be seen that the elevator deflection angle is adjusted quickly to achieve active control of elastic vibration. Figure 6It can be seen that the pitch angles measured by the four pitch angle sensors are quite different, including both the rigid body pitch angle and the elastic pitch angle. Figure 7 It can be seen that the tracking error of the rigid body pitch angle gradually approaches 0, achieving accurate tracking of the rigid body pitch angle. Figures 8a to 8c It can be seen that the generalized coordinate derivatives of the first three elastic modes gradually decay, achieving rapid suppression of the elastic mode.

[0123] It should be understood that the various forms of processes shown above can be used to reorder, add or delete steps. For example, the steps described in the present invention can be executed in parallel, sequentially or in different orders, as long as the desired results of the technical solution disclosed in the present invention can be achieved, and this document does not limit this.

[0124] The above specific implementations do not constitute a limitation on the protection scope of the present invention. It should be understood by those skilled in the art that various modifications, combinations, sub-combinations and substitutions can be made according to design requirements and other factors. Any modification, equivalent substitution and improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. An aerodynamic servoelastic active control system for an air-breathing hypersonic vehicle, characterized in that: The system comprises: an elastic frequency estimator, a state disturbance joint estimator and an attitude controller, wherein the elastic frequency estimator is used to obtain the elastic frequency estimation value of the controlled object by using an elastic frequency estimation model based on the pitch angle information sent by the controlled object, and send the elastic frequency estimation value to the state disturbance joint estimator; the state disturbance joint estimator is used to estimate the rigid body state, the elastic body state, the disturbance factor of the rigid body state, the disturbance factor of the elastic body state and the pitch angle information by using the state disturbance joint estimation model based on the received pitch angle information, the elevator deflection angle and the elastic frequency estimation value, obtain the corresponding state estimation result, the disturbance estimation result and the pitch angle estimation result, and send the state estimation result and the disturbance estimation result to the attitude controller; the attitude controller is used to obtain the elevator deflection angle of the controlled object by using the attitude control model based on the received rigid body pitch angle control value, the state estimation result and the disturbance estimation result, and send the elevator deflection angle to the state disturbance joint estimator and the controlled object.

2. The system according to claim 1, characterized in that in, The elastic frequency estimation model meets the following conditions: θ(t)=θ(t-ΔT)+(1-σ)R -1 (t)ψ θ (t)e(t); R(t)=σR(t-ΔT)+(1-σ)[ψ θ (t)ψ θ (t) T +βI]; Among them, v(t) is the input of the elastic frequency estimation model at time t, e(t) is the output of the elastic frequency estimation model at time t, and are the pitch angles measured by the d-th pitch angle sensor and the k-th pitch angle sensor, respectively. The values ​​of k and d are 1 to m, and k≠d. m is the number of pitch angle sensors. represents the vibration mode φ of the i-th order elastic mode corresponding to the k-th pitch angle sensor i (p k ), represents the vibration mode φi(p d ), θ i (t) is The estimation function of is the estimated value of the elastic frequency of the i-th elastic mode among the first three elastic modes used at time t, i ranges from 1 to 3, ε is the preset adjustment parameter, z -1 and z -2 is the delay operator, β is the step adjustment factor; θ(t) is the estimation function of the elastic frequency at time t, is the estimated value of the elastic frequency of the first three elastic modes used at time t, △T is the sampling time; σ is the forgetting factor; R(t) is the covariance matrix at time t, ψ θ (t) is the gradient of e(t) with respect to θ(t), and I is the identity matrix; The state disturbance joint estimation model meets the following conditions: in, is the state of the model jointly estimated for the state disturbance, for The derivative of is the rigid body pitch angle The estimated value of is the estimated value of the rigid body pitch angular velocity Q, is the generalized coordinate η of the i-th elastic mode i The estimated value of is the generalized coordinate η of the i-th elastic mode i The derivative of The estimated value of is the perturbation factor d of the derivative of Q Q The estimated value of for The perturbation factor of the derivative of The estimated value of is the pitch angle estimation result, is the pitch angle measured by the hth pitch angle sensor The estimated value of h ranges from 1 to m; δ e is the elevator deflection angle, y is the pitch angle information, E is the gain of the state disturbance joint estimation model, B1 and C are the state matrix, input matrix and output matrix respectively; where, C=[C00 m×4 ], in, q is the flight pressure, S is the reference area of ​​the aircraft, is the longitudinal reference length of the aircraft, z T is the coupling coefficient from thrust T to pitch moment M, is η i The influence coefficient on the pitch matrix M, is η i Influence coefficient on thrust T, I yy is the moment of inertia of the aircraft; Represents η i The generalized force N for the i-th elastic mode i The influence coefficient of ξ i is the damping of the i-th elastic mode; is the elevator and canard rudder deflection angle δ e The influence coefficient on the pitching moment M, is the canard rudder deflection angle δ c The influence coefficient on M, is δ e The influence coefficient on lift L, is δ c The influence coefficient on lift L; Denotes δ e To N i The influence coefficient of Denotes δ c To N i The influence coefficient of p k represents the distance between the kth pitch angle sensor and the head of the aircraft, represents the vibration mode φ of the i-th order elastic mode corresponding to the k-th pitch angle sensor i (p k )’s first-order derivative; The posture control model meets the following conditions: Among them, k p , k d and To preset control parameters, It is the rigid body pitch angle control value.

3. The system according to claim 1, characterized in that m satisfies the following conditions: Among them, λ j for The eigenvalue of , j ranges from 1 to 12, and rank() represents the rank of the matrix; The position vector of m sensors p=[p1 p2Λp m ]The following conditions are met: cond(O j ) represents the matrix O corresponding to the jth eigenvalue j The condition number of the pitch angle sensor is shown in Figure 2, J(p) is the position objective function, and P is the feasible interval of the pitch angle sensor position.

4. The system according to claim 1, characterized in that The controlled object is an air-breathing hypersonic aircraft or a proxy model of an air-breathing hypersonic aircraft.

5. The system according to claim 4, characterized in that The proxy model of the air-breathing hypersonic vehicle meets the following conditions: d c =rδ e ; Where V represents the speed of the aircraft, is the first-order derivative of V, γ represents the ballistic inclination angle of the aircraft, is the first-order derivative of γ, h represents the height, is the first-order derivative of h, α represents the angle of attack of the aircraft, is the first-order derivative of α, is η i The first derivative of is η i The second derivative of i represents the true elastic frequency of the i-th elastic mode, G represents the mass of the aircraft; D represents the resistance of the aircraft, φ represents the fuel equivalence ratio of the aircraft, C T η represents the coupling coefficient vector from thrust T to generalized coordinate η, C T，φ (·), C L (·), C D (·), C M (·), C T (·) are aerodynamic parameters, N i b Indicates b to N i The influence coefficient of b belongs to (α 2 , α, δ e , δ c ,η),N i 0 N i The corresponding constant, Indicates the pitch angle measured by the pitch angle sensor, represents the elastic pitch angle, represents the vibration mode of the i-th elastic mode φ i The first derivative of (p), where p represents the distance from the pitch angle sensor to the head of the aircraft.

6. An active control method for aerodynamic servoelasticity of an air-breathing hypersonic vehicle, characterized in that: The steps include: S100, in response to receiving the rigid body pitch angle control value at the current moment, the attitude controller obtains the elevator deflection angle of the controlled object using the attitude control model, and sends it to the state disturbance joint estimator and the controlled object; S200, the controlled object performs a corresponding operation based on the received elevator deflection angle, obtains pitch angle information after performing the corresponding operation, and sends the information to the elastic frequency estimator; S300, based on the received pitch angle information, the elastic frequency estimator obtains an elastic frequency estimation value of the controlled object using an elastic frequency estimation model, and sends the elastic frequency estimation value to the state disturbance joint estimator; S400, based on the received pitch angle information, elevator deflection angle and elastic frequency estimation value, estimate the rigid body state, elastic body state, the disturbance factor of the rigid body state, the disturbance factor of the elastic body state and the pitch angle information using a state-disturbance joint estimation model to obtain corresponding state estimation results, disturbance estimation results and pitch angle estimation results, and send the state estimation results and disturbance estimation results to the attitude controller; Execute S100.

7. The method according to claim 6, characterized in that in, The elastic frequency estimation model meets the following conditions: θ(t)=θ(t-ΔT)+(1-σ)R -1 (t)ψ θ (t)e(t); R(t)=σR(t-ΔT)+(1-σ)[ψ θ (t)ψ θ (t) T +βI]; Among them, v(t) is the input of the elastic frequency estimation model at time t, e(t) is the output of the elastic frequency estimation model at time t, and are the pitch angles measured by the d-th pitch angle sensor and the k-th pitch angle sensor, respectively. The values ​​of k and d are 1 to m, and k≠d. m is the number of pitch angle sensors. represents the vibration mode φ of the i-th order elastic mode corresponding to the k-th pitch angle sensor i (p k ), represents the vibration mode φi(p d ), θ i (t) is The estimation function of is the estimated value of the elastic frequency of the i-th elastic mode among the first three elastic modes used at time t, i ranges from 1 to 3, ε is the preset adjustment parameter, z -1 and z -2 is the delay operator, β is the step adjustment factor; θ(t) is the estimation function of the elastic frequency at time t, is the estimated value of the elastic frequency of the first three elastic modes used at time t, △T is the sampling time; σ is the forgetting factor; R(t) is the covariance matrix at time t, ψ θ (t) is the gradient of e(t) with respect to θ(t), and I is the identity matrix; The state disturbance joint estimation model meets the following conditions: in, is the state of the model jointly estimated for the state disturbance, for The derivative of is the rigid body pitch angle The estimated value of is the estimated value of the rigid body pitch angular velocity Q, is the generalized coordinate η of the i-th elastic mode i The estimated value of is the generalized coordinate η of the i-th elastic mode i The derivative of The estimated value of is the perturbation factor d of the derivative of Q Q The estimated value of for The perturbation factor of the derivative of The estimated value of is the pitch angle estimation result, is the pitch angle measured by the hth pitch angle sensor The estimated value of e is the elevator deflection angle, y is the pitch angle information, E is the gain of the state disturbance joint estimation model, B1 and C are the state matrix, input matrix and output matrix respectively; where, C=[C00 m×4 ], in, q is the flight pressure, S is the reference area of ​​the aircraft, is the longitudinal reference length of the aircraft, z T is the coupling coefficient from thrust T to pitch moment M, is η i The influence coefficient on the pitch matrix M, is η i Influence coefficient on thrust T, I yy is the moment of inertia of the aircraft; Represents η i The generalized force N for the i-th elastic mode i The influence coefficient of ξ i is the damping of the i-th elastic mode; is the elevator deflection angle δ e The influence coefficient on the pitching moment M, is the canard rudder deflection angle δ c The influence coefficient on M, is δ e The influence coefficient on lift L, is δ c The influence coefficient on lift L; Denotes δ e To N i The influence coefficient of Denotes δ c To N i The influence coefficient of p k represents the distance between the kth pitch angle sensor and the head of the aircraft, represents the vibration mode φ of the i-th order elastic mode corresponding to the k-th pitch angle sensor i (p k )’s first-order derivative; The posture control model meets the following conditions: Among them, k p , k d and To preset control parameters, It is the rigid body pitch angle control value.

8. The method according to claim 7, characterized in that m satisfies the following conditions: Among them, λ j for The eigenvalue of , j ranges from 1 to 12, and rank() represents the rank of the matrix; The position vector of m sensors p=[p1 p2Λp m ]The following conditions are met: cond(O j ) represents the matrix O corresponding to the jth eigenvalue j The condition number of the pitch angle sensor is shown in Figure 2, J(p) is the position objective function, and P is the feasible interval of the pitch angle sensor position.

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