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

Abstract

The application provides a kind of air-breathing hypersonic vehicle aerodynamic servo elasticity active control system and method, comprising: elasticity frequency estimator obtains the elasticity frequency estimation value of controlled object based on the pitch angle information sent by controlled object, and sends to state disturbance joint estimator;State disturbance joint estimator obtains corresponding state estimation result, disturbance estimation result and pitch angle estimation result based on the pitch angle information received, elevator deflection angle and elasticity frequency estimation value;Attitude controller obtains the elevator deflection angle of controlled object based on the rigid body pitch angle control value received, state estimation result and disturbance estimation result, and sends to state disturbance joint estimator and controlled object.The application can control the stability of controlled object.

Description

Technical Field

[0001] This invention relates to the field of air-breathing hypersonic vehicle control, and in particular to an aerodynamic servo-elastic active control system and method for air-breathing hypersonic vehicles. Background Technology

[0002] Currently, most research on the control of elastically air-breathing hypersonic vehicles assumes that rigid body motion information is measurable and only considers aeroelastic effects. However, sensor measurement signals simultaneously contain rigid body motion information and elastic vibration information, and the rigid body motion information cannot be directly obtained. Therefore, in addition to aeroelasticity, controller design must also consider the servoelasticity caused by elastic vibration feedback in the measurement signals, resulting in a complex aero-servoelastic effect. Aero-servoelastic control methods for elastically air-breathing hypersonic vehicles mainly include robust control, amplitude stabilization, phase stabilization, and active control. Robust control characterizes the desired elastic mode suppression performance through a closed-loop performance transfer function and obtains the controller by solving an optimization problem. Amplitude stabilization passively suppresses elastic modes by designing notch filters. Compared to amplitude stabilization, phase stabilization can extend the rigid body control bandwidth and obtain satisfactory stability margin. Phase stabilization, by optimizing the positions of sensors and actuators, enables self-stabilization of elastic modes by directly feeding the measurement signals back to the flight control system; it is a semi-active control method with fixed elastic state feedback gain. In order to suppress multiple elastic modes simultaneously and quickly, it is necessary to study active control methods that can design the feedback gain of each elastic state independently. Summary of the Invention

[0003] To address the aforementioned technical problems, the technical solution adopted by this invention is as follows:

[0004] According to a first aspect of the present invention, an aerodynamic servo-elastic active control system for an air-breathing hypersonic vehicle is provided. The system includes: an elastic frequency estimator, a state disturbance joint estimator, and an attitude controller. The elastic frequency estimator is used to obtain an estimated elastic frequency value of the controlled object based on pitch angle information sent by the controlled object using an elastic frequency estimation model, and sends this estimated 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 based on received pitch angle information, elevator deflection angle, and the estimated elastic frequency value, using the state disturbance joint estimator model to obtain corresponding state estimation results, disturbance estimation results, and pitch angle estimation results, and sends 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 based on received rigid body pitch angle control values, state estimation results, and disturbance estimation results using an attitude control model, and sends this angle to the state disturbance joint estimator and the controlled object.

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

[0006] S100, in response to receiving the rigid body pitch angle control value at the current moment, the attitude controller uses the attitude control model to obtain the elevator deflection angle of the controlled object 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 after performing the corresponding operation, obtains pitch angle information and sends it to the elastic frequency estimator.

[0008] S300, based on the received pitch angle information, the elastic frequency estimator uses the elastic frequency estimation model to obtain the elastic frequency estimate of the controlled object and sends it to the joint state disturbance estimator.

[0009] S400: Based on the received pitch angle information, elevator deflection angle, and elastic frequency estimate, the state disturbance joint estimation model is used to estimate the rigid body state, elastic body state, disturbance factor of the rigid body state, disturbance factor of the elastic body state, and pitch angle information, to obtain the corresponding state estimation results, disturbance estimation results, and pitch angle estimation results, and the state estimation results and disturbance estimation results are sent to the attitude controller; S100 is executed.

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

[0011] The aerodynamic servo-elastic active control system and method for air-breathing hypersonic vehicles provided in this invention have a lower dependence on model accuracy compared to robust control. Specifically, a disturbance factor is introduced into the nominal observation model. This disturbance factor is the model error, including parameter uncertainties not considered during the establishment of the nominal observation model, unmodeled dynamics, and external disturbances. This invention first estimates the disturbance factor using a joint state disturbance estimator, and then compensates for the disturbance factor in the attitude controller to eliminate its influence. Therefore, even when the model has strong uncertainties, the designed control system can still ensure system stability. In addition, the elastic frequency can be estimated in real time, reducing the conservatism of the controller. Compared to amplitude stability, this invention uses an active controller and does not use a notch filter, avoiding the phase lag introduced by the notch filter, thus not affecting the dynamic performance and stability margin of rigid body control. Compared to phase stability, this invention can simultaneously suppress multiple elastic modes through the control parameters in the attitude controller, and by flexibly designing the closed-loop damping of each elastic mode, the elastic modes can achieve the desired suppression speed.

[0012] It should be understood that the description in this section is not intended to identify key or essential features of the embodiments of the present invention, nor is it intended to limit the scope of the invention. Other features of the invention will become readily apparent from the following description. Attached Figure Description

[0013] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0014] Figure 1 This is a structural block diagram of the aerodynamic servo-elastic active control system for an air-breathing hypersonic vehicle provided in an embodiment of the present invention;

[0015] Figure 2 A flowchart of an active aerodynamic servo-elastic control method for an air-breathing hypersonic vehicle provided in an embodiment of the present invention;

[0016] Figure 3 This is a schematic diagram of the frequency estimation results from the frequency estimator.

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

[0018] Figure 5 A schematic diagram of the control variables calculated for the attitude controller;

[0019] Figure 6 A schematic diagram showing the measured values ​​of each sensor obtained after the control input is fed into the surrogate model;

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

[0021] Figures 8a to 8c This is a schematic diagram of the generalized coordinate derivatives of the first three elastic modes. Detailed Implementation

[0022] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort 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 commonly understood by one of ordinary skill in the art to which this invention pertains. The terminology used herein in the description of this invention is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0024] It should be noted that some exemplary embodiments are described as processes or methods depicted as flowcharts. Although the flowcharts describe the steps as sequential processes, many of these steps can be performed in parallel, concurrently, or simultaneously. Furthermore, the order of the steps can be rearranged. A process can be terminated when its operation is complete, but it may also have additional steps not included in the figures. A process can correspond to a method, function, procedure, subroutine, subroutine, etc.

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

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

[0027] In this embodiment of the invention, the controlled object can be a real air-breathing hypersonic vehicle or a proxy model of an air-breathing hypersonic vehicle. That is, the active control system provided in this embodiment of the invention can control the attitude of the actual flying vehicle and can also be used for attitude control simulation analysis of the vehicle.

[0028] Furthermore, in this embodiment of the invention, the surrogate model of the elastic air-breathing hypersonic vehicle is modeled as a free beam, such that the rigid body mode and the elastic mode satisfy orthogonality, that is, the rigid body mode and the elastic mode only influence each other through the interaction of forces and moments. To improve the non-minimum phase characteristics of the trajectory inclination angle, the surrogate model also introduces canard deflection. Specifically, the surrogate model of the air-breathing hypersonic vehicle satisfies the following conditions:

[0029]

[0030] δ c =rδ e (3);

[0031]

[0032]

[0033] Where V represents the speed of the aircraft. Let γ be the first derivative of V, and let γ represent the trajectory inclination angle of the aircraft. Let γ be the first derivative of γ, and h represent the height. Let α be the first derivative of h, and let α represent the angle of attack of the aircraft. Let α be the first derivative, and Q represent the pitch rate of the aircraft. The first derivative of Q, Indicates the pitch angle of the aircraft. for The first derivative; η i Let represent the generalized coordinates of the i-th elastic mode out of the first three elastic modes, where i takes values ​​from 1 to 3. For η i The first derivative, For η i The second derivative; ξ i ω represents the damping of the i-th elastic mode. i The i-th elastic mode represents the true elastic frequency. G represents the mass of the aircraft; I represents the mass of the i-th elastic mode. yy The moment of inertia of the aircraft is represented by T; the thrust acting on the aircraft is represented by D; the drag acting on the aircraft is represented by L; the lift acting on the aircraft is represented by M; and the pitching moment acting on the aircraft is represented by N. i Let represent the generalized force of the i-th elastic mode, q represent the flight kinetic pressure of the aircraft, and S represent the reference area of ​​the aircraft. δ represents the longitudinal reference length of the aircraft. e For elevator deflection angle, δ c The canard deflection angle is φ, where φ represents the aircraft's fuel equivalence ratio, and C... T η z represents the coupling coefficient vector from the thrust T to the generalized coordinate η.T C represents the coupling coefficient from thrust T to pitching moment M. T，φ (·), C L (·), C D (·), C M (·), C T (·) are all aerodynamic parameters, N i b Indicates b with respect to N i The influence coefficient, b belongs to (α) 2 , α, δ e δ c ,η),N i 0 N represents i The corresponding constant, δ e The influence coefficient on L δ c The influence coefficient on L; This indicates the pitch angle measured by the pitch angle sensor. Indicates the elastic pitch angle, The mode shape φ represents the i-th elastic mode. i The first derivative of (p), where p represents the distance between the pitch angle sensor and the nose of the aircraft.

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

[0035] The expression for the aerodynamic coefficient is as follows:

[0036]

[0037] The expression in the expression This represents the influence coefficient of * on ·.

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

[0039] In this embodiment of the invention, only the 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, for example, 1.5.

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

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

[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] Where v(t) is the input of the elastic frequency estimation model at time t, and e(t) is the output of the elastic frequency estimation model at time t. and These are the pitch angles measured by the d-th and k-th pitch angle sensors, respectively. The values ​​of k and d range from 1 to m, and k ≠ d. m is the number of pitch angle sensors. The mode shape φ represents the i-th elastic mode corresponding to the k-th pitch angle sensor. i (p k The first derivative of ) The mode shape φ represents the i-th elastic mode corresponding to the d-th pitch angle sensor. i (p d The first derivative of ) means that the k-th pitch angle sensor and the d-th pitch angle sensor can be any two sensors selected from the m pitch angle sensors.

[0046] Where, θ i (t) is The estimated function, This is the estimated elastic frequency of the i-th elastic mode out of the first three elastic modes used, at time t, where i takes values ​​from 1 to 3, ε is a preset adjustment parameter, and z... -1 and z -2 Here, β is the delay operator, β is the step size adjustment factor, and θ(t) is the estimation function of the elastic frequency at time t. Here, ΔT represents the estimated elastic frequencies of the first three elastic modes at time t, where ΔT is the sampling time; σ is the forgetting factor; R(t) is the covariance matrix at time t, and ψ... θ (t) is the gradient of e(t) with respect to θ(t), and I is the identity matrix;

[0047] In this embodiment of the invention, β is used to influence the stability and convergence speed of parameter updates, and can be a preset fixed value. The larger ε is, the slower the convergence speed of the frequency estimator, but the higher the estimation accuracy. -1 and z -2 These represent delaying the signal by 1 sample time and 2 sample times, respectively. θ i The initial value of ψ(t) can be 0. θ (t) can be obtained using existing methods.

[0048] In practical applications, the elastic frequency ω i The elastic frequency varies considerably during flight, and using a fixed value would fail to achieve adaptive suppression of elastic vibration. Therefore, this invention uses a frequency estimator to estimate the elastic frequency in real time, thereby enabling adaptive suppression of elastic vibration.

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

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

[0051] In this embodiment of the invention, the nominal observation model serves as the basis for selecting the number and location of pitch angle sensors, and is also the foundation for designing the joint state perturbation estimator. In this embodiment, based on the aforementioned surrogate model, the following linear small perturbation model for attitude angles is constructed:

[0052]

[0053] Where, x s The state represents the linear small perturbation model of the attitude angle. u2 represents the total perturbation in the linear small perturbation model of the attitude angle. y represents the pitch angle information of the linear small perturbation model strategy for attitude angles. Let d be the pitch angle measured by the h-th pitch angle sensor. Q The perturbation factor of the derivative of Q. for The disturbance factor of the derivative means that the control system model in this embodiment of the invention has strong uncertainty, which can be eliminated by estimating the total disturbance. A0 is the state matrix of the linear small disturbance model of attitude angle, B1 and B2 are the input state matrices of the linear small disturbance model of attitude angle, and C0 is the output matrix of the linear small disturbance model of attitude angle, where:

[0054]

[0055] The expression for each element in each matrix is ​​as follows:

[0056]

[0057] in, Indicates η i The generalized force N for the i-th elastic mode i Influence coefficient, ξ i For the damping of the i-th elastic mode, p k This represents the distance between the k-th pitch angle sensor and the nose of the aircraft. The mode shape φ represents the i-th elastic mode corresponding to the k-th pitch angle sensor. i (p k The first derivative of ).

[0058] A complete small attitude angle perturbation model needs to include the dynamics of α; however, it has been proven that α is unobservable. Therefore, the small attitude angle perturbation model does not include the dynamics of α, but incorporates the effect of α on Q into d. Q middle.

[0059] Since the total disturbance vector u2 is unknown, it is extended to the system state in order to estimate u2 simultaneously. The extended system state is as follows: The expanded system model is re-represented as

[0060]

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

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

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

[0064]

[0065] in, The state is the state of the joint estimation model for state perturbations. for The derivative, pitch angle of rigid body The estimated value, Let Q be an estimated value of the rigid body pitch angular velocity. Let η be the generalized coordinate of the i-th elastic mode. i The estimated value, The generalized coordinates η of the first i-th elastic modes i derivative The estimated value, The perturbation factor d of the derivative of Q Q The estimated value, for The perturbation factor of the derivative The estimated value, For the pitch angle estimation results, The pitch angle measured by the h-th pitch angle sensor The estimated value of h, where h ranges from 1 to m; δ e y represents the elevator deflection angle, and y represents the pitch angle information. E represents the gain of the joint estimation model for state perturbations, 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 this embodiment of the invention, the selection of the number and location configuration of the pitch angle sensors need to ensure that the joint estimation model of state perturbation is observable, that is, it can estimate x through y. a To make an estimate, in order to analyze observability, it is first necessary to determine... The eigenvalues. After calculation, It has 6 eigenvalues ​​of 0 and 6 non-zero eigenvalues, namely

[0067] Furthermore, the necessary and sufficient condition for the observability of the joint estimation model of state perturbations is:

[0068]

[0069] Where, λ j for The eigenvalues ​​of the matrix, where j ranges from 1 to 12, and rank() represents the rank of the matrix. m×4 Let m be an m-row, m-column matrix containing only zeros. That is, m must satisfy the above condition.

[0070] The following discussion will cover two scenarios:

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

[0072]

[0073] Where M1 = [m1, m2, ..., m 12 To calculate the rank of M1, we perform a series of elementary column transformations on M1. First, we adjust the order of the first 8 columns of M1, resulting in:

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

[0075] Then, using the last four columns of M1, the elements in rows 2, 4, 6, and 8 of A0 are all set to 0, meaning A0 becomes A0 = [0, ... 8×4 A1], where At this point, M1 has been transformed into At this point, rank(M) = rank(C1) + rank([A1 B2]) = rank(C1) + 8. To satisfy the above conditions, C1 needs to be full rank. Therefore, m ≥ 4, meaning the minimum number of pitch angle sensors is 4. Furthermore, the positional configuration of the pitch angle sensors needs to meet the full rank requirement.

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

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

[0078] Then, after a series of elementary row operations, we can obtain Clearly, rank(N) = 8, which satisfies the observability requirement.

[0079] In summary, full rank of column C1 is a necessary and sufficient condition for ensuring the observability of the joint estimation model of state perturbations.

[0080] In this embodiment of the invention, the positions of each sensor are further determined based on maximizing the key elastic mode observability matrix. Let p = [p1 p2Λp] m [] represents the position vectors of m (m≥4) sensors, and calculates the matrix corresponding to the j-th eigenvalue. condition number cond(O) j The higher the condition number, the more likely the matrix is ​​to become non-full rank, and the more easily the observability condition is violated. Therefore, the choice of p can be achieved by solving the following optimization problem:

[0081]

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

[0083] Since the problem of maximizing and minimizing the condition number is non-convex, numerical optimization algorithms are generally used to solve for p. This invention uses a genetic algorithm to solve for p. First, a population is initialized, with each individual in the population corresponding 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. This process is repeated iteratively until the stopping condition is met. At this point, the number and location of the sensors are determined, and the output matrix C in the joint estimation model of state perturbation can be obtained.

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

[0085]

[0086] Where, k p k d and For preset control parameters, This is the control value for the rigid body pitch angle.

[0087] For ease of analysis, it is assumed that there is a frequency band separation between the rigid body dynamics and elastic body dynamics of an air-breathing hypersonic vehicle, δ e,1 and δ e,2 These are used for the control of rigid bodies and elastic bodies, respectively. δ e,1 Bring into The dynamic equations can be obtained as follows:

[0088]

[0089] Assuming the estimation errors of the state and disturbances approach 0, k p >0, k d >0, then It will asymptotically converge to 0. Similarly, δ e,2 Substitute into η i The dynamic equations, neglecting the coupling effects between different elastic modes and the state estimation errors, yield the following:

[0090]

[0091] It is evident that introducing feedback from the generalized coordinate derivative of the elastic mode increases the damping of elastic vibration, thereby accelerating the rapid decay of elastic vibration until... It approaches 0.

[0092] In this embodiment of the invention, the parameters in the controller It can change the suppression speed of the i-th elastic mode, and can be designed individually for different elastic modes according to task requirements.

[0093] Based on the same inventive concept, embodiments of the present invention provide a method for aerodynamic servo-elastic active control of an air-breathing hypersonic vehicle, such as... Figure 2 As shown, the steps may include the following:

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

[0095] In this embodiment of the invention, the controlled object can be a real air-breathing hypersonic vehicle or a proxy model of an air-breathing hypersonic vehicle. That is, the active control system provided in this embodiment of the invention can control the attitude of the actual flying vehicle and can also be used for attitude control simulation analysis of the vehicle.

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

[0097] In this embodiment of the invention, when the controlled object is an aircraft, since the canard deflection angle can be calculated based on the elevator deflection angle, the aircraft's elevator deflection angle can be controlled to match the received elevator deflection angle by the aircraft's elevator deflection angle control mechanism, and the aircraft's canard deflection angle can be controlled to match the received canard deflection angle by the aircraft's canard deflection angle control mechanism. Then, the aircraft's pitch angle sensor measures the pitch angle at this time and sends it to the elastic frequency estimator. When the controlled object is a proxy model of the aircraft, the canard deflection angle can be calculated based on the received elevator deflection angle. Then, the elevator deflection angle, canard deflection angle, and fuel equivalence ratio are used as inputs to the proxy model. The proxy model obtains the corresponding pitch angle information and various states in the model. The initial values ​​of each state are determined based on the initial state of the controlled object, for example, the initial flight altitude can be 10km.

[0098] S300, based on the received pitch angle information, the elastic frequency estimator uses the elastic frequency estimation model to obtain the elastic frequency estimate of the controlled object and sends it to the joint state disturbance estimator.

[0099] S400: Based on the received pitch angle information, elevator deflection angle, and elastic frequency estimate, the state disturbance joint estimation model is used to estimate the rigid body state, elastic body state, disturbance factor of the rigid body state, disturbance factor of the elastic body state, and pitch angle information, to obtain the corresponding state estimation results, disturbance estimation results, and pitch angle estimation results, and the state estimation results and disturbance estimation results are sent to the attitude controller; S100 is executed.

[0100] The control method of this embodiment can be implemented by the aforementioned control system. Therefore, for the specific implementation of this method, please refer to the embodiment shown in the aforementioned system, which will not be elaborated further.

[0101] (Example)

[0102] (1) Some parameters in the control-related elastic surrogate model of a certain air-breathing hypersonic vehicle are as follows. Based on the following parameters, all parameters in the joint estimation model of state disturbance 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] This invention employs four pitch angle sensors, with the sensor positions arranged to satisfy observability conditions. The mode shape derivatives of the elastic modes at each sensor position are as follows:

[0109]

[0110] (2) Based on the above parameters, the parameters of the joint estimation model of state disturbance 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) Taking into account the influence 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 With an eigenvalue of -30, the design parameter E of the joint state disturbance estimator can be calculated as follows:

[0119]

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

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

[0122] Based on the above configuration parameters, the frequency estimation result of the frequency estimator can be as follows: 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 results of the joint state perturbation estimator are as follows: Figures 4a to 4e As shown, according to Figures 4a to 4e It can be seen that the error between the estimation result and the actual result of the joint state disturbance estimator is very small, almost equal. The control quantity calculated by the attitude controller can be as follows: Figure 5 As shown, the measured values ​​of each sensor obtained after the control input is fed into the surrogate model can be as follows: Figure 6 As shown, the tracking error of the rigid body pitch angle is as follows: Figure 7 As shown, the generalized coordinate derivatives of the first three elastic modes are as follows: Figures 8a to 8c As shown. According to Figure 5 It can be seen that the elevator deflection angle is rapidly adjusted to achieve active control of elastic vibration. According to... Figure 6It can be seen that the pitch angles measured by the four pitch angle sensors differ significantly, encompassing both rigid body pitch angles and elastic pitch angles. According to... Figure 7 It can be seen that the tracking error of the rigid body pitch angle gradually approaches 0, thus achieving accurate tracking of the rigid body pitch angle. According to... Figures 8a to 8c It can be seen that the generalized coordinate derivatives of the first three elastic modes gradually decay, thus achieving rapid suppression of the elastic modes.

[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 this invention can be executed in parallel, sequentially, or in different orders, as long as the desired result of the technical solution disclosed in this invention can be achieved, and this is not limited herein.

[0124] The specific embodiments described above do not constitute a limitation on the scope of protection of this invention. Those skilled in the art should understand that various modifications, combinations, sub-combinations, and substitutions can be made according to design requirements and other factors. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of this invention should be included within the scope of protection of this invention.

Claims

1. A pneumatic servo-elastic active control system for an air-breathing hypersonic vehicle, characterized in that, The system includes: an elastic frequency estimator, a state disturbance joint estimator, and an attitude controller. The elastic frequency estimator, based on pitch angle information sent by the controlled object, uses an elastic frequency estimation model to obtain an estimated elastic frequency value for the controlled object and sends it to the state disturbance joint estimator. The state disturbance joint estimator, based on received pitch angle information, elevator deflection angle, and the estimated elastic frequency value, uses the state disturbance joint estimator to estimate the rigid body state, elastic body state, disturbance factor of the rigid body state, disturbance factor of the elastic body state, and pitch angle information, obtaining corresponding state estimation results, disturbance estimation results, and pitch angle estimation results, and sends the state estimation results and disturbance estimation results to the attitude controller. The attitude controller, based on received rigid body pitch angle control values, state estimation results, and disturbance estimation results, uses an attitude control model to obtain the elevator deflection angle of the controlled object and sends it to the state disturbance joint estimator and the controlled object. The joint estimation model for state perturbations satisfies the following conditions: in, The state is the state of the joint estimation model for state perturbations. for The derivative of pitch angle of rigid body The estimated value, Let Q be an estimated value of the rigid body pitch angular velocity. Let η be the generalized coordinate of the i-th elastic mode. i The estimated value, Let η be the generalized coordinate of the i-th elastic mode. i derivative The estimated value, The perturbation factor d of the derivative of Q Q The estimated value, for The perturbation factor of the derivative The estimated value, For the pitch angle estimation results, The pitch angle measured by the h-th pitch angle sensor The estimated value of h, where h ranges from 1 to m; δ e y represents the elevator deflection angle, and y represents the pitch angle information. E represents the gain of the joint estimation model for state perturbations. B1 and C are the state matrix, input matrix, and output matrix, respectively; where... C = [C00] m×4 ], in, q represents the flight pressure, and S represents the reference area of ​​the aircraft. z is the longitudinal reference length of the aircraft. T Let T be the coupling coefficient from thrust T to pitching moment M. For η i Influence coefficient on pitch matrix M For η i The influence coefficient of thrust T, I yy The moment of inertia of the aircraft; Indicates η i The generalized force N for the i-th elastic mode i Influence coefficient, ξ i The damping is for the i-th elastic mode; Elevator canard deflection angle δ e Influence coefficient on pitching moment M For the canard deflection angle δ c The influence coefficient on M For δ e The influence coefficient on lift L For δ c The influence coefficient on lift L; δ e For N i Influence coefficient, δ c For N i The influence coefficient; p k This represents the distance between the k-th pitch angle sensor and the nose of the aircraft. The mode shape φ represents the i-th elastic mode corresponding to the k-th pitch angle sensor. i (p k The first derivative of ) The attitude control model satisfies the following conditions: Where, k p k d and For preset control parameters, This is the control value for the rigid body pitch angle.

2. The system according to claim 1, characterized in that, in, The elastic frequency estimation model satisfies 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]; Where v(t) is the input of the elastic frequency estimation model at time t, and e(t) is the output of the elastic frequency estimation model at time t. and These are the pitch angles measured by the d-th and k-th pitch angle sensors, respectively. The values ​​of k and d range from 1 to m, and k ≠ d. m is the number of pitch angle sensors. The mode shape φ represents the i-th elastic mode corresponding to the k-th pitch angle sensor. i (p k The first derivative of ) The mode shape φ represents the i-th elastic mode corresponding to the d-th pitch angle sensor. i (p d The first derivative of θ i (t) is The estimated function, This is the estimated elastic frequency of the i-th elastic mode out of the first three elastic modes used, at time t, where i takes values ​​from 1 to 3, ε is a preset adjustment parameter, and z... -1 and z -2 Here, β is the delay operator, β is the step size adjustment factor, and θ(t) is the estimation function of the elastic frequency at time t. Here, ΔT represents the estimated elastic frequencies of the first three elastic modes at time t, where ΔT is the sampling time; σ is the forgetting factor; R(t) is the covariance matrix at time t, and ψ... θ (t) is the gradient of e(t) with respect to θ(t), and I is the identity matrix.

3. The system according to claim 2, characterized in that, m satisfies the following condition: Where, λ j for The eigenvalues ​​of , j takes values ​​from 1 to 12, and rank() represents the rank of the matrix; The position vector p of m sensors = [p1 p2Λp] m The following conditions must be met: cond(O j ) represents the matrix O corresponding to the j-th eigenvalue. j The condition number is J(p), which is the position objective function, and P represents 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 vehicle or a proxy model of an air-breathing hypersonic vehicle.

5. The system according to claim 4, characterized in that, The surrogate model of the air-breathing hypersonic vehicle satisfies the following conditions: d c =rδ e ; Where V represents the speed of the aircraft. Let γ be the first derivative of V, and let γ represent the trajectory inclination angle of the aircraft. Let γ be the first derivative of γ, and h represent the height. Let α be the first derivative of h, and let α represent the angle of attack of the aircraft. The first derivative of α For η i The first derivative, For η i The second derivative; ω i Let G represent the true elastic frequency of the i-th elastic mode, G represent the mass of the aircraft, D represent the drag experienced by the aircraft, φ represent the fuel equivalence ratio of the aircraft, and C represent the mass of the aircraft. T η C represents the coupling coefficient vector from thrust T to generalized coordinate η. T，φ (·), C L (·), C D (·), C M (·), C T (·) are all aerodynamic parameters, N i b Indicates b with respect to N i The influence coefficient, b belongs to (α) 2 , α, δ e δ c ,η),N i 0 N represents i The corresponding constant, This indicates the pitch angle measured by the pitch angle sensor. Indicates the elastic pitch angle. The mode shape φ represents the i-th elastic mode. i The first derivative of (p), where p represents the distance between the pitch sensor and the nose of the aircraft. for The first derivative.

6. A method for aerodynamic servo-elastic active control of an air-breathing hypersonic vehicle, characterized in that, Includes the following steps: S100, in response to receiving the rigid body pitch angle control value at the current moment, the attitude controller uses the attitude control model to obtain the elevator deflection angle of the controlled object 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, and after performing the corresponding operation, obtains pitch angle information and sends it to the elastic frequency estimator. S300, based on the received pitch angle information, the elastic frequency estimator uses the elastic frequency estimation model to obtain the elastic frequency estimate of the controlled object and sends it to the state disturbance joint estimator. S400: Based on the received pitch angle information, elevator deflection angle, and elastic frequency estimate, the state disturbance joint estimation model is used to estimate the rigid body state, elastic body state, disturbance factor of the rigid body state, disturbance factor of the elastic body state, and pitch angle information, to obtain the corresponding state estimation results, disturbance estimation results, and pitch angle estimation results, and the state estimation results and disturbance estimation results are sent to the attitude controller; S100 is executed. The joint estimation model for state perturbations satisfies the following conditions: in, The state is the state of the joint estimation model for state perturbations. for The derivative of pitch angle of rigid body The estimated value, Let Q be an estimated value of the rigid body pitch angular velocity. Let η be the generalized coordinate of the i-th elastic mode. i The estimated value, Let η be the generalized coordinate of the i-th elastic mode. i derivative The estimated value, The perturbation factor d of the derivative of Q Q The estimated value, for The perturbation factor of the derivative The estimated value, For the pitch angle estimation results, The pitch angle measured by the h-th pitch angle sensor The estimated value, δ e y represents the elevator deflection angle, and y represents the pitch angle information. E represents the gain of the joint estimation model for state perturbations. B1 and C are the state matrix, input matrix, and output matrix, respectively; where... C = [C00] m×4 ], in, q represents the flight pressure, and S represents the reference area of ​​the aircraft. z is the longitudinal reference length of the aircraft. T Let T be the coupling coefficient from thrust T to pitching moment M. For η i Influence coefficient on pitch matrix M For η i The influence coefficient of thrust T, I yy The moment of inertia of the aircraft; Indicates η i The generalized force N for the i-th elastic mode i Influence coefficient, ξ i The damping is for the i-th elastic mode; Elevator deflection angle δ e Influence coefficient on pitching moment M For the canard deflection angle δ c The influence coefficient on M For δ e The influence coefficient on lift L For δ c The influence coefficient on lift L; δ e For N i Influence coefficient, δ c For N i The influence coefficient; p k This represents the distance between the k-th pitch angle sensor and the nose of the aircraft. The mode shape φ represents the i-th elastic mode corresponding to the k-th pitch angle sensor. i (p k The first derivative of ) The attitude control model satisfies the following conditions: Where, k p k d and For preset control parameters, This is the control value for the rigid body pitch angle.

7. The method according to claim 6, characterized in that, in, The elastic frequency estimation model satisfies 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]; Where v(t) is the input of the elastic frequency estimation model at time t, and e(t) is the output of the elastic frequency estimation model at time t. and These are the pitch angles measured by the d-th and k-th pitch angle sensors, respectively. The values ​​of k and d range from 1 to m, and k ≠ d. m is the number of pitch angle sensors. The mode shape φ represents the i-th elastic mode corresponding to the k-th pitch angle sensor. i (p k The first derivative of ) The mode shape φ represents the i-th elastic mode corresponding to the d-th pitch angle sensor. i (p d The first derivative of θ i (t) is The estimated function, This is the estimated elastic frequency of the i-th elastic mode out of the first three elastic modes used, at time t, where i takes values ​​from 1 to 3, ε is a preset adjustment parameter, and z... -1 and z -2 Here, β is the delay operator, β is the step size adjustment factor, and θ(t) is the estimation function of the elastic frequency at time t. Here, ΔT represents the estimated elastic frequencies of the first three elastic modes at time t, where ΔT is the sampling time; σ is the forgetting factor; R(t) is the covariance matrix at time t, and ψ... θ (t) is the gradient of e(t) with respect to θ(t), and I is the identity matrix.

8. The method according to claim 7, characterized in that, m satisfies the following condition: Where, λ j for The eigenvalues ​​of , j takes values ​​from 1 to 12, and rank() represents the rank of the matrix; The position vector p of m sensors = [p1 p2Λp] m The following conditions must be met: cond(O j ) represents the matrix O corresponding to the j-th eigenvalue. j The condition number is J(p), which is the position objective function, and P represents the feasible interval of the pitch angle sensor position.

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