An enhanced anti-interference control system and method for an aircraft fairing separation process

Abstract

The invention discloses an enhanced anti-interference control system and method for an aircraft fairing separation process, comprising modeling additional interference to the fairing separation process and aircraft attitude dynamics; simplifying the attitude dynamics model to obtain a control-oriented system model; designing an enhanced anti-interference control law for the control-oriented system model based on an adaptive backstepping method, wherein adaptive neural network technology is used to compensate for unknown parts in the model, and a tanh function boundary estimator is used to process inner-loop uncertainty and outer-loop interference in the model; completing stability analysis of the designed enhanced anti-interference control law based on Lyapunov stability theory, and verifying the effectiveness of the invention through simulation examples, thereby achieving the goal of achieving an expected value of the aircraft attitude under the additional interference generated in the fairing separation process and under the conditions of unknown parts and uncertainty in the model.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft, and in particular relates to an enhanced anti-interference control system and method for an aircraft fairing separation process. Background Art

[0002] During the separation of the fairing, the aircraft is mainly disturbed by the following three situations: first, the impact interference of the explosive bolts; second, the sudden change of the flow field caused by the change of the aerodynamic shape and overall parameters of the body after the fairing is detached; third, the flow field interference caused by the forward and lateral movement of the fairing. Accordingly, the difficulty of suppressing the interference of the fairing separation on the aircraft to control the attitude movement of the aircraft can be summarized as follows: first, the aircraft has different aerodynamic characteristics before and after separation, and it is necessary to consider the interference caused by the explosive bolts and the change of aerodynamic shape during the separation of the fairing, and it is difficult for traditional flight control methods to achieve effective interference suppression; second, during separation, the fairing is in front and the aircraft is behind, and in order to ensure safety, the fairing needs to complete forward and lateral movement. Under high-speed and high-dynamic pressure conditions, these movements of the fairing will impose shock wave interference on the aircraft body, further increasing the difficulty of interference suppression of the aircraft. Traditional aircraft control methods cannot suppress the interference of the aircraft during the separation of the fairing, thereby affecting the attitude movement of the aircraft. Therefore, it is of great theoretical and practical significance to study the enhanced anti-interference control method of the aircraft during the separation of the fairing. Summary of the invention

[0003] The purpose of the present invention is to overcome the shortcomings of the above-mentioned prior art and provide an enhanced anti-interference control system and method for the separation process of the aircraft fairing, so as to solve the problem that the prior art aircraft control method cannot eliminate or compensate for the interference to the aircraft during the separation process of the fairing. The present invention studies the problem of aircraft attitude control during the separation process of the fairing, based on the analysis of the multi-source uncertainty generated by the separation of the fairing, and in view of the characteristics and difficulties of the sudden change of aerodynamic shape and shock wave interference in the control process, studies a new control method with better interference suppression capability and control effect.

[0004] In order to achieve the above object, the present invention adopts the following technical solutions:

[0005] An enhanced anti-interference control method for an aircraft fairing separation process comprises the following steps:

[0006] Step 1, establishing a mathematical model of the additional interference to the aircraft when the fairing is separated from the aircraft, and establishing an attitude dynamics model of the aircraft based on the mathematical model;

[0007] Step 2, simplifying the attitude dynamics model of the aircraft into a control-oriented system model, wherein the system model includes an inner loop subsystem and an outer loop subsystem;

[0008] Step 3, establishing a dynamic equation of the inner loop subsystem tracking error and a dynamic equation of the outer loop subsystem tracking error, and establishing an inner loop virtual control law and an outer loop control law based on the two dynamic equations;

[0009] When the fairing separates from the aircraft, the aircraft executes the control law, and when the inner loop tracking error tends to zero and the outer loop tracking error tends to zero, the attitude of the aircraft tends to be stable.

[0010] A further improvement of the present invention is:

[0011] Preferably, in step 1, the mathematical model of the additional interference is:

[0012]

[0013] in, are the segmented moment values ​​applied by the fairing to the aircraft in the longitudinal and lateral directions, n is the total number of steps, t i The time when each separation stage starts.

[0014] Preferably, in step 1, the attitude dynamics model of the aircraft is a system model of the pitch and yaw channels of the aircraft.

[0015] Preferably, the system model of the pitch and yaw channels of the aircraft is:

[0016]

[0017] Among them, α is the angle of attack of the aircraft, and β is the sideslip angle of the aircraft; represents the derivative of the aerodynamic moment coefficient with respect to the angle of attack and the sideslip angle; δ z ,δ y They represent elevator deviation and rudder deviation respectively; Expresses the derivatives of the aerodynamic moment coefficient with respect to the elevator and rudder deflections; represents the derivative of the moment coefficient with respect to the angle of attack and the angle of sideslip; Represents the derivative of the moment coefficient with respect to the rudder deviation; ω z ,ω y are the longitudinal and lateral angular velocities of the aircraft respectively; It represents the derivative of the torque coefficient with respect to the dimensionless coefficient of angular velocity; Q, S, L are the aerodynamic pressure, reference area and reference length of the aircraft respectively; m, V are the mass and velocity of the aircraft respectively; F z ,F y are the longitudinal and lateral forces of the aircraft respectively; J z ,J y are the longitudinal and lateral moments of inertia of the aircraft respectively; f z (α,β,ω z ,ωy ),f y (α,β,ω z ,ω y ) is the unmodeled part of the aircraft system model; d ZLZ,z (t),d ZLZ,y (t) is the longitudinal and lateral disturbances experienced by the aircraft.

[0018] Preferably, in step 2, the inner ring subsystem and the outer ring subsystem They are:

[0019]

[0020] in,

[0021]

[0022] Among them, α is the angle of attack of the aircraft, β is the sideslip angle of the aircraft; ω z ,ω y are the longitudinal and lateral angular velocities of the aircraft respectively; δ z ,δ y represents the elevator rudder deviation; Q, S, L are respectively the aerodynamic pressure, reference area and reference length of the aircraft; m, V are respectively the mass and speed of the aircraft; represents the derivative of the aerodynamic moment coefficient with respect to the angle of attack and the sideslip angle; It expresses the derivative of the aerodynamic moment coefficient with respect to the elevator and rudder deflections.

[0023] Preferably, in step 3, the inner loop control law x 2c for:

[0024]

[0025] The outer loop control law u is:

[0026]

[0027] Among them, e1 is the inner loop tracking error, e2 is the outer loop tracking error, k1, k2>0, both are feedback gain coefficients to be designed, y d =[α d β d ] T , α d ,β d Respectively represent the desired angle of attack and sideslip angle signals of the aircraft, D1=sup||d1||, D2=sup||d2+ε Θ ||, ε Θ is the estimated error of the neural network fit, is the estimated value of the weight, and ε1,ε2>0 are small constants.

[0028] Preferably, the dynamic equation of the inner loop tracking error and the dynamic equation of the outer loop tracking error are respectively:

[0029] e1=x1-y d ,e2=x2-x 2c (6)

[0030] In step 3, the uncertainty of the inner loop subsystem and the interference of the outer loop subsystem are estimated and compensated online through the tanh function boundary estimator through the following formula (9):

[0031]

[0032] Preferably, in step 3, the stability analysis of the inner loop virtual control rate and the outer loop control law is performed by using Lyapunov stability theory.

[0033] An enhanced anti-interference control system for an aircraft fairing separation process, comprising:

[0034] A dynamics module, used to establish a mathematical model of additional interference imposed by the fairing on the aircraft when the fairing is separated from the aircraft, and to establish an attitude dynamics model of the aircraft based on the mathematical model;

[0035] A system model module, used to simplify the attitude dynamics model of the aircraft into a control-oriented system model, wherein the system model includes an inner loop subsystem and an outer loop subsystem;

[0036] A control law module is used to establish a dynamic equation of an inner loop tracking error and a dynamic equation of an outer loop tracking error, and to establish an inner loop virtual control rate and an outer loop control law based on the two dynamic equations;

[0037] When the fairing separates from the aircraft, the aircraft executes the control law, and the separation process becomes stable when the inner loop tracking error tends to zero and the outer loop tracking error tends to zero.

[0038] Preferably, it also includes:

[0039] The analysis module is used to perform stability analysis on the inner and outer control loops using Lyapunov stability theory.

[0040] Compared with the prior art, the present invention has the following beneficial effects:

[0041] The invention discloses an enhanced anti-interference control method for an aircraft fairing separation process, comprising modeling additional interference to the fairing separation process and aircraft attitude dynamics; simplifying the attitude dynamics model to obtain a control-oriented system model; designing an enhanced anti-interference control law for the control-oriented system model based on an adaptive backstepping method, wherein adaptive neural network technology is used to compensate for unknown parts in the model, and a tanh function boundary estimator is used to process inner-loop uncertainty and outer-loop interference in the model; completing stability analysis of the designed enhanced anti-interference control law based on Lyapunov stability theory, and verifying the effectiveness of the invention through simulation examples, thereby achieving the goal of achieving an expected value for the aircraft attitude under the additional interference generated in the fairing separation process and under the conditions of unknown parts and uncertainty in the model. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] Figure 1 This is the interference form diagram generated during the hood separation process;

[0043] Figure 2 is a control flow chart of the present invention;

[0044] Figure 3 The following are the angle of attack tracking effects in three different situations;

[0045] Figure 4 The following are the side slip angle tracking effects in three different situations;

[0046] Figure 5 The diagrams of elevator deviation control input signals under three different situations are shown;

[0047] Figure 6 The rudder deviation control input signal diagrams under three different situations;

[0048] Figure 7 The compensation curve diagram of inner loop interference under three different situations;

[0049] Figure 8 The compensation curves of the outer loop interference under three different situations. DETAILED DESCRIPTION

[0050] The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments:

[0051] An aircraft hood enhanced anti-interference control method comprises the following steps:

[0052] Step 1: Establish a mathematical model of the additional interference to the aircraft during the fairing separation process, and construct an attitude dynamics model of the aircraft during the fairing separation process.

[0053] During the fairing separation process, the interference to the aircraft exists in the pitch channel and the yaw channel: the impact interference of the explosive bolts mainly acts on the pitch channel, and the shape parameters and flow field interference act on the pitch and yaw channels at the same time. Therefore, only the attitude control problem of the pitch and yaw channels of the aircraft is considered. Analyzing the characteristics of the additional interference to the aircraft during the fairing separation process, it can be known that the interference generated during the hood separation process is as follows: Figure 1 As shown, based on Figure 1 , the disturbance (load) caused by the separation of the nose cover on the attitude motion of the hypersonic vehicle can be expressed as the following formula (1):

[0054]

[0055] Among them, d ZLZ,y (t),d ZLZ,z (t) is a step function with non-equal intervals, which is specifically expressed as follows:

[0056]

[0057] in, are the segmented moment values ​​in pitch and yaw directions applied by the fairing to the hypersonic aircraft, n is the total number of steps, t i The time when each separation stage starts.

[0058] The system model for constructing the pitch channel and yaw channel of the aircraft during the fairing separation process is:

[0059]

[0060] Among them, α is the angle of attack of the aircraft, and β is the sideslip angle of the aircraft; represents the derivative of the aerodynamic moment coefficient with respect to the angle of attack and the sideslip angle; δ z ,δ y They represent elevator deviation and rudder deviation respectively; Expresses the derivatives of the aerodynamic moment coefficient with respect to the elevator and rudder deflections; represents the derivative of the moment coefficient with respect to the angle of attack and the angle of sideslip; Represents the derivative of the moment coefficient with respect to the rudder deviation; ω z ,ω y are the longitudinal and lateral angular velocities of the aircraft respectively; It represents the derivative of the torque coefficient with respect to the dimensionless coefficient of angular velocity; Q, S, L are the aerodynamic pressure, reference area and reference length of the aircraft respectively; m, V are the mass and velocity of the aircraft respectively; F z ,F y are the longitudinal and lateral forces of the aircraft respectively; J z ,J y are the longitudinal and lateral moments of inertia of the aircraft respectively; fz (α,β,ω z ,ω y ),f y (α,β,ω z ,ω y ) is the unmodeled part of the aircraft system model; d ZLZ,z (t),d ZLZ,y (t) is the disturbance experienced by the aircraft in the longitudinal and lateral directions.

[0061] Step 2: Simplify the attitude dynamics model of the aircraft during the fairing separation process, that is, formula (3), to obtain a control-oriented system model. In order to facilitate the design of the control law, it is necessary to simplify the attitude dynamics model of the aircraft. First, define the state vector x1(t) = [αβ] T ,x2(t)=[ω z ω y ] T and control input vector u(t) = [δ z ,δ y ] T , then based on the defined state vector, the attitude dynamics model (3) of the aircraft is vectorized, and finally the control-oriented system model is obtained, as shown in the following formula (4). Based on the design idea of ​​the backstepping method, the control-oriented system model is regarded as a system consisting of an inner loop subsystem and the outer ring subsystem It consists of two subsystems.

[0062]

[0063] Among them, x2(t), d1(t), x1(t), B, u(t), d2(t)

[0064]

[0065] Step 3: Design an enhanced anti-interference control law for the fairing separation process of the aircraft. Based on the idea of ​​backstepping, in order to make the attitude of the aircraft change according to the given command, that is, x1 can track the reference signal y d , need x2 according to x 2c Therefore, it is necessary to design the control law u to make x2 meet the design requirements. Based on the above analysis, we first define the inner loop tracking error and the outer loop tracking error e1 = x1-y d ,e2=x2-x 2c Based on the aircraft control-oriented system model (4) and the defined tracking error, the dynamic equations of the inner and outer loop tracking errors are established:

[0066]

[0067] Where x1(t) = [αβ] T ,x2(t)=[ω z ω y ] T ,y d =[α d β d ] T ,x 2c (t) = [ω zc ω yc ] T , α d ,β d Respectively represent the desired angle of attack and sideslip angle signals of the aircraft, ω zc ,ω yc Indicates the desired longitudinal and lateral angular velocity of the aircraft.

[0068] For the unknown part of the system, f(x1(t),x2(t)) in the specific formula (4) is compensated using adaptive neural network technology:

[0069]

[0070] Among them, Φ(x1(t),x2(t)) uses Gaussian function, is the estimated value of f(x1(t),x2(t)), ε Θ is the estimation error, Θ T is the weight of the neural network, is the estimated value of Θ, and its update rate is

[0071]

[0072] Γ Θ ,λ Θ >0 is the adaptive parameter to be designed.

[0073] Next, define D1 = sup||d1||, D2 = sup||d2+ε Θ ||, use the tanh function boundary estimator to estimate and compensate the uncertainty of the inner loop subsystem and the interference of the outer loop subsystem online, define is the estimated value of D1 and D2, and its update rate is designed to be:

[0074]

[0075] Among them, λ1,λ2>0 are parameters to be designed, and ε1,ε2>0 are small constants.

[0076] Based on the design idea of ​​backstepping method, the inner loop virtual control law (x 2c ) so that x1 tracks to y d, e1 approaches 0, and the outer loop control law u makes x2 track to x 2c , making e2 approach to 0. When this situation is reached, the attitude system of the aircraft tends to be stable.

[0077]

[0078]

[0079] Among them, k1, k2>0 are the feedback gain coefficients to be designed.

[0080] The control block diagram of the system is as follows Figure 2 shown.

[0081] Step 4: Based on the Lyapunov stability theory, complete the stability analysis of the enhanced anti-interference control law for fairing separation. Select the Lyapunov function:

[0082]

[0083] Taking the derivative of both sides of the equation:

[0084]

[0085] Substituting the dynamic equation (6) of the tracking error, the control laws (10), (11) and the adaptive laws (8), (9) into the above equation (13), the simplified equation is:

[0086]

[0087] in,

[0088]

[0089] According to Lyapunov's stability theorem, the system eventually converges to a bounded state and is definitely stable.

[0090] Example

[0091] The designed fairing separation enhanced anti-interference control law is simulated and verified using MATLAB. The simulation example verification is completed under three situations, and the model parameters of the three different working conditions are selected as shown in Table 1 below:

[0092] Table 1

[0093]

[0094] Figure 3 Figure 2 is a diagram showing the angle of attack tracking effect under three different situations in the present invention. Figure 4 The following are the side slip angle tracking effect diagrams under three different situations in the present invention. Figure 5 are the elevator deviation control input signals under three different situations in the present invention, Figure 6 are the rudder deviation control input signals under three different situations in the present invention, Figure 7 are the compensation curves of the inner loop interference under three different situations in the present invention, Figure 8 The compensation curves of the outer loop interference under three different situations in the present invention are shown. The above simulation result diagrams show that the present invention can effectively converge the system.

[0095] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. An enhanced anti-interference control method for an aircraft fairing separation process, characterized in that: The following steps are involved: Step 1, establishing a mathematical model of the additional interference to the aircraft when the fairing is separated from the aircraft, and establishing an attitude dynamics model of the aircraft based on the mathematical model; In step 1, the attitude dynamics model of the aircraft is a system model of the pitch and yaw channels of the aircraft; The system model of the aircraft pitch and yaw channels is: （3） in, is the angle of attack of the aircraft, is the sideslip angle of the aircraft; represents the derivative of the aerodynamic moment coefficient with respect to the angle of attack and the sideslip angle; They represent elevator deviation and rudder deviation respectively; Expresses the derivatives of the aerodynamic moment coefficient with respect to the elevator and rudder deflections; represents the derivative of the moment coefficient with respect to the angle of attack and the angle of sideslip; It expresses the derivative of the moment coefficient with respect to the rudder deviation; are the longitudinal and lateral angular velocities of the aircraft respectively; represents the derivative of the torque coefficient with respect to the dimensionless coefficient of angular velocity; are the aerodynamic pressure, reference area and reference length of the aircraft respectively; are the mass and speed of the aircraft respectively; are the longitudinal and lateral forces of the aircraft respectively; are the longitudinal and lateral moments of inertia of the aircraft respectively; It is the unmodeled part of the aircraft system model; is the longitudinal and lateral disturbances to which the aircraft is subjected; Step 2, simplifying the attitude dynamics model of the aircraft into a control-oriented system model, wherein the system model includes an inner loop subsystem and an outer loop subsystem; Step 3, establishing a dynamic equation of the inner loop subsystem tracking error and a dynamic equation of the outer loop subsystem tracking error, and establishing an inner loop virtual control law and an outer loop control law based on the two dynamic equations; When the fairing separates from the aircraft, the aircraft executes the control law, and when the inner loop tracking error tends to zero and the outer loop tracking error tends to zero, the attitude of the aircraft tends to be stable.

2. The enhanced anti-interference control method for the separation process of an aircraft fairing according to claim 1, characterized in that: In step 1, the mathematical model of the additional interference is: （2） in, are the segmented moment values ​​applied by the fairing to the aircraft in the longitudinal and lateral directions, is the total number of steps, The time when each separation stage starts.

3. The enhanced anti-interference control method for the separation process of an aircraft fairing according to claim 1, characterized in that: In step 2, the inner ring subsystem and the outer ring subsystem They are: （4） in, （5） in, is the angle of attack of the aircraft, is the sideslip angle of the aircraft; are the longitudinal and lateral angular velocities of the aircraft respectively; Indicates elevator rudder deviation; are the aerodynamic pressure, reference area and reference length of the aircraft respectively; are the mass and speed of the aircraft respectively; represents the derivative of the aerodynamic moment coefficient with respect to the angle of attack and the sideslip angle; It expresses the derivative of the aerodynamic moment coefficient with respect to the elevator and rudder deflections.

4. The enhanced anti-interference control method for the separation process of an aircraft fairing according to claim 1, characterized in that: In step 3, the inner loop virtual control law for: （10） The outer loop control law u for: （11） in, e 1 is the inner loop tracking error, e 2 is the outer loop tracking error, , are the feedback gain coefficients to be designed, , They represent the desired angle of attack and sideslip angle signals of the aircraft respectively, , is the estimated error of the neural network fit, is the estimated value of the weight, is a small constant.

5. The enhanced anti-interference control method for the separation process of an aircraft fairing according to claim 4, characterized in that: The dynamic equations of the inner loop tracking error and the outer loop tracking error are respectively: （6） In step 3, the uncertainty of the inner loop subsystem and the interference of the outer loop subsystem are estimated and compensated online through the tanh function boundary estimator through the following formula (9): （9）。 6. The enhanced anti-interference control method for the separation process of an aircraft fairing according to claim 1, characterized in that: In step 3, the stability of the inner loop virtual control law and the outer loop control law is analyzed by Lyapunov stability theory.

7. An enhanced anti-interference control system for realizing the aircraft fairing separation process as claimed in claim 1, characterized in that: include: A dynamics module, used to establish a mathematical model of additional interference imposed by the fairing on the aircraft when the fairing is separated from the aircraft, and to establish an attitude dynamics model of the aircraft based on the mathematical model; A system model module, used to simplify the attitude dynamics model of the aircraft into a control-oriented system model, wherein the system model includes an inner loop subsystem and an outer loop subsystem; A control law module is used to establish a dynamic equation of an inner loop tracking error and a dynamic equation of an outer loop tracking error, and to establish an inner loop virtual control law and an outer loop control law based on the two dynamic equations; When the fairing separates from the aircraft, the aircraft executes the control law, and the separation process becomes stable when the inner loop tracking error tends to zero and the outer loop tracking error tends to zero.

8. The enhanced anti-interference control system for the separation process of an aircraft fairing according to claim 7, characterized in that: Also includes: The analysis module is used to perform stability analysis on the inner loop virtual control law and the outer loop control law through Lyapunov stability theory.

Images