Hypersonic flight vehicle fixed time control method and system based on event triggering

Abstract

The invention relates to a hypersonic flight vehicle fixed time control method and system based on event triggering. The method comprises the following steps: constructing a control-oriented attitude system model of a hypersonic flight vehicle; designing a switching dynamic event triggering strategy based on the attitude system model, and establishing an internal dynamic variable in the switching dynamic event triggering strategy; constructing a fixed time sliding mode controller by adopting a fixed time control method according to the angle tracking error and the reaching law; constructing an event triggering fixed time sliding mode controller according to the switching dynamic event triggering strategy and the fixed time sliding mode controller; error information of the aircraft is obtained in real time, the error information is input into an event to trigger a fixed time sliding mode controller for calculation, and attitude control input is obtained; and adjusting and controlling the attitude angle of the aircraft through attitude control input. Through the method of the invention, complex and lengthy hierarchical design can be avoided, the resource overhead of an aircraft flight control system is reduced while the robustness is ensured, and the method has practical engineering significance.

Description

Technical Field

[0001] The present application relates to the field of aircraft control technology, and in particular to a method and system for fixed-time control of a hypersonic aircraft based on event triggering. Background Art

[0002] Hypersonic vehicles are an important aerospace technology for entering near-space. Since the United States conducted multiple hypersonic flight tests in the last century, major countries around the world have begun to develop hypersonic vehicles, and hypersonic vehicles have become an important aerospace technology that is indispensable for competition among major countries. In recent years, shape-shifting vehicles have attracted widespread attention from aerospace research institutions due to their advantages in improving aerodynamic characteristics.

[0003] Hypersonic morphing aircraft can continuously morph to obtain aerodynamics that are more suitable for the current environment and mission, achieve better flight performance, and even realize multi-mission design. Morphing technology not only brings better aerodynamic characteristics to hypersonic aircraft, but also brings many challenges. The drastic changes in aerodynamic characteristics require the design of a control system with strong robustness, good adaptability and low resource consumption for hypersonic aircraft. Summary of the invention

[0004] Based on this, it is necessary to provide an event-triggered hypersonic aircraft fixed-time control method and system that can save aircraft system resources and has high tracking performance and low resource consumption characteristics in order to address the above technical problems.

[0005] A method for controlling a hypersonic vehicle at a fixed time based on event triggering, the method comprising:

[0006] Construct a control-oriented attitude system model for hypersonic vehicles;

[0007] Based on the posture system model, a switching dynamic event triggering strategy is designed, and an internal dynamic variable whose convergence characteristics match the sliding mode variable is established in the switching dynamic event triggering strategy;

[0008] According to the angle tracking error and reaching law, a fixed-time sliding mode controller is constructed by using the fixed-time control method.

[0009] Constructing an event-triggered fixed-time sliding mode controller according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller;

[0010] Acquire error information of the aircraft in real time, input the error information into the event-triggered fixed-time sliding mode controller for calculation, and obtain attitude control input;

[0011] The attitude angle of the aircraft is adjusted and controlled through the attitude control input.

[0012] In one embodiment, constructing a control-oriented attitude system model of a hypersonic vehicle includes:

[0013] Based on the original nonlinear model of the hypersonic vehicle, a control-oriented attitude system model of the hypersonic vehicle is constructed.

[0014] In one embodiment, the original nonlinear model of the hypersonic vehicle is expressed by a nonlinear function, and the nonlinear function expression is:

[0015] First, define the scalar σ, which is expressed as:

[0016]

[0017] Construct a nonlinear function expression:

[0018]

[0019]

[0020] Where, σ=[σ1 σ2 … σ n ] Τ Represents the element as σ j ,j=1,...,n vector; l1,l2 represent the designable positive parameters respectively; and Ψ(σ1,σ2; l1,l2) denote pre-designed functions; ||·|| denotes the 2-norm.

[0021] In one embodiment, the control-oriented attitude system model of the hypersonic vehicle is expressed as:

[0022]

[0023] in,

[0024]

[0025]

[0026]

[0027]

[0028] In the formula, x1=[α β γ V ] represents the attitude angle; x2=[ω z ω y ω x ] represents the attitude angular velocity; u=[δ z δ y δ x ] represents the rudder deflection angle; di , i = 1, 2 represents the aircraft modeling error and external interference, and the complex terms in the attitude system motion equation and the additional moment caused by the deformation are combined into d i ; L, D, N represent lift, drag, and lateral force, respectively; V, m, g represent flight speed, hypersonic vehicle mass, and gravitational acceleration, respectively; I x ,I y ,I z represents the three-axis moment of inertia; θ represents the track angle; q represents the dynamic pressure; They are respectively expressed as the pitch moment coefficient related to the pitch rudder deflection angle, the yaw moment coefficient term related to the yaw rudder deflection angle, and the roll moment coefficient term related to the roll rudder deflection angle; S0 is the reference area; α is the vector composed of the angle of attack α=[1 α α 3 ] Τ , β mx =β; is the pitching moment coefficient term related to α; is the yaw moment coefficient term related to the sideslip angle; is the rolling moment coefficient term related to the roll angle.

[0029] In one embodiment, the switching dynamic event triggering strategy is expressed as:

[0030]

[0031]

[0032] In the formula, u j (t) represents the jth component of the control quantity u(t); υ j Indicates the event trigger controller that will be designed; and Indicates the start and end time of the time interval of the jth channel under the event trigger mechanism; λ i ,i=1,2,3,δ are coefficients;η j Represents an internal dynamic variable; z u,j represents the difference between the event-triggered controller and the ideal control input; δ represents the constant coefficient used to switch the trigger condition.

[0033] In one embodiment, the internal dynamic variable expression is:

[0034]

[0035] In the formula, c r , r=1,2,3,4 are coefficients; Represents a nonlinear function with independent variable y; y = [s Τ ηΤ b] Τ represents the vector required in the design process; s represents the sliding surface to be designed; η is the vector form of the internal dynamic variable.

[0036] In one embodiment, a fixed-time sliding mode controller is constructed using a fixed-time control method according to an angle tracking error and a reaching law, including:

[0037] The angular tracking error is expressed as:

[0038] z=x1-x c =[z1 z2 z3] Τ ;

[0039] According to the angle tracking error z and the derivative function of the error variable Design sliding surface s = [s1 s2 s3] Τ , the expression is:

[0040]

[0041] The sliding surface is made to converge within a fixed time through the reaching law. The reaching law expression is:

[0042]

[0043] Then, the fixed-time sliding mode controller function is designed as:

[0044]

[0045] In the formula, k r , r = 2, 3, 4 and μ j ,j=1,2,3,4 are coefficients.

[0046] In one embodiment, the event-triggered fixed-time sliding mode controller is expressed as:

[0047]

[0048] Where υ0(t) represents the control input; x c =[x 1c x 2c x 3c ] Τ Indicates angle command; s = [s1 s2 s3] Τ represents the sliding surface; k and ε are coefficients; Indicates a positive designable coefficient.

[0049] In one embodiment, the gesture control input is represented as:

[0050]

[0051] Where: Ψ(y,s; k3,k4) represents the function with independent variables y and s; k0 and k1 are coefficients, k1>L1+D2, k3=c3, k4=c4; L1 and D2 represent the upper bounds of the absolute values ​​of the interference and the reciprocal of the interference respectively.

[0052] A hypersonic vehicle fixed time control system based on event triggering, the system comprising:

[0053] The attitude system model building module is used to build the attitude system model for control of hypersonic aircraft;

[0054] A switching dynamic event trigger strategy design module is used to design a switching dynamic event trigger strategy based on the posture system model, and establish an internal dynamic variable whose convergence characteristics match the sliding mode variable in the switching dynamic event trigger strategy;

[0055] A fixed-time sliding mode controller building module is used to build a fixed-time sliding mode controller using a fixed-time control method according to an angle tracking error and a reaching law;

[0056] An event-triggered fixed-time sliding mode controller construction module is used to construct an event-triggered fixed-time sliding mode controller according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller;

[0057] An attitude control input calculation module is used to obtain error information of the aircraft in real time, input the error information into the event-triggered fixed-time sliding mode controller for calculation, and obtain attitude control input;

[0058] The attitude adjustment control module is used to adjust and control the attitude angle of the aircraft through the attitude control input.

[0059] The above-mentioned event-triggered fixed-time control method and system for a hypersonic aircraft constructs a control-oriented attitude system model for a hypersonic aircraft; then, based on the attitude system model, a switching dynamic event triggering strategy is designed, and internal dynamic variables whose convergence characteristics match the sliding mode variables are established in the switching dynamic event triggering strategy; a fixed-time sliding mode controller is constructed using a fixed-time control method according to the angle tracking error and the convergence law; an event-triggered fixed-time sliding mode controller is constructed according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller; the error information of the aircraft is obtained in real time, and the error information is input into the event-triggered fixed-time sliding mode controller for calculation to obtain the attitude control input; and the attitude angle of the aircraft is adjusted and controlled through the attitude control input.

[0060] On the one hand, this application designs a switching dynamic event trigger strategy to ensure robustness while reducing the resource overhead of the aircraft flight control system, thereby achieving high-performance and low-consumption attitude control, which has practical engineering significance; on the other hand, by constructing a fixed-time sliding mode controller, the ideal control quantity is directly obtained, avoiding complex and lengthy hierarchical design, and can achieve more ideal control performance when applied to hypersonic aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0061] Figure 1 A schematic diagram of the steps of a method for controlling a hypersonic aircraft at a fixed time based on event triggering in one embodiment;

[0062] Figure 2 It is a schematic diagram of a flow chart of fixed-time control of a hypersonic aircraft based on event triggering in one embodiment;

[0063] Figure 3 A schematic diagram of a tracking curve of an attitude angle to a command in a simulation example 1 in an embodiment;

[0064] Figure 4 A schematic diagram of a curve showing a change in angular velocity of a simulation example 1 in an embodiment;

[0065] Figure 5 A schematic diagram of a curve showing a change in rudder deflection angle and deformation rate of a simulation example 1 in an embodiment;

[0066] Figure 6 A schematic diagram of changes in the control quantity update time and interval of a simulation example 1 in an embodiment;

[0067] Figure 7 A schematic diagram of a comparison curve of attitude angle response of a simulation example 2 in an embodiment;

[0068] Figure 8 A schematic diagram of a curve showing a change in angular velocity of a simulation example 2 in an embodiment;

[0069] Fig. 9 A schematic diagram of the change of the pitch rudder deflection angle in simulation example 2 and the change of the control amount update time and update interval under different methods in an embodiment;

[0070] Fig.10 A schematic diagram of the change of the yaw rudder angle in simulation example 2 and the change of the control amount update time and update interval under different methods in an embodiment;

[0071] Fig.11 A schematic diagram of the change of the roll rudder angle in simulation example 2 and the change of the control amount update time and update interval under different methods in an embodiment;

[0072] Fig.121 is a structural block diagram of a fixed-time control system for a hypersonic aircraft based on event triggering in one embodiment. DETAILED DESCRIPTION

[0073] In order to make the purpose, technical solution and advantages of the present application more clearly understood, the present application is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and are not used to limit the present application.

[0074] In the process of implementing the technical solution of the present application, in order to make the hypersonic aircraft have strong robustness while not increasing the resource computing burden of the system, the inventor constructed an event-triggered fixed-time sliding mode controller by designing a switching dynamic event triggering strategy and a fixed-time sliding mode controller. Based on the aircraft error information obtained in real time, the attitude control input can be directly obtained through the event-triggered fixed-time sliding mode controller. After the control input is input to the hypersonic aircraft, the aircraft attitude can track the set angular command within a fixed time, thereby ensuring that the aircraft attitude is quickly stabilized and saving system resources by reducing the controller update frequency.

[0075] Among them, the fixed-time control method can make the closed-loop system error converge to the far point within a fixed time. Compared with the traditional finite-time control technology, the convergence time of the closed-loop system using the fixed-time control method is independent of the initial conditions of the system, which improves the influence of the larger system initial value on the convergence time. By combining the fixed-time control method with the multivariable sliding mode control technology, the ideal control quantity can be directly obtained, avoiding complex and lengthy hierarchical design, and can achieve more ideal control performance when applied to hypersonic aircraft.

[0076] The switching dynamic event trigger design is designed to save aircraft system resources. Modern robust controllers are often based on real-time updates of control quantities. The more robust the controller, the higher the control quantity update frequency required. However, hypersonic aircraft are not systems with unlimited resources, such as computing power, communication resources, and actuator deflection limits, which are all limited. Therefore, using an event trigger mechanism to ensure robustness while reducing the resource overhead of the aircraft flight control system to achieve high-performance and low-consumption attitude control has practical engineering significance.

[0077] The following will describe the implementation of the present invention in detail with reference to the accompanying drawings in the embodiment diagram of the present invention.

[0078] In one embodiment, Figure 1 and Figure 2 As shown, a method for controlling a hypersonic vehicle at a fixed time based on event triggering is provided, comprising the following steps:

[0079] Step 102, constructing a control-oriented attitude system model of a hypersonic vehicle.

[0080] Specifically, based on the original nonlinear model of the hypersonic aircraft, a control-oriented attitude system model of the hypersonic aircraft was constructed.

[0081] More specifically, the attitude system motion equation of the hypersonic vehicle attitude dynamics model is expressed as:

[0082]

[0083] Among them, α, β, γ V Respectively represent the angle of attack, sideslip angle, and roll angle; ω x ,ω y ,ω z M represents the roll, yaw and pitch angular velocity respectively; L, D and N represent the lift, drag and lateral force respectively; V, m and g represent the flight speed, mass of the hypersonic vehicle and gravitational acceleration respectively. x ,M y ,M z Respectively represent the rolling, yaw and pitching moments; M sx ,M sy ,M sz They represent the additional moments of the three-axis deformation caused by the corresponding deformation; I x ,I y ,I z represents the three-axis moment of inertia; θ represents the track angle.

[0084] The original nonlinear model of the hypersonic vehicle is expressed by a nonlinear function, and the nonlinear function expression is:

[0085] First, define the scalar σ, which is expressed as:

[0086]

[0087] Construct a nonlinear function expression:

[0088]

[0089]

[0090] Where, σ=[σ1 σ2 … σ n ] Τ Represents the element as σ j ,j=1,...,n vector; l1,l2 represent the designable positive parameters respectively; and Ψ(σ1,σ2; l1,l2) denote pre-designed functions; ||·|| denotes the 2-norm.

[0091] Therefore, the control-oriented attitude system model of a hypersonic vehicle is expressed as:

[0092]

[0093] in,

[0094]

[0095]

[0096]

[0097]

[0098] In the formula, x1=[α β γ V ] represents the attitude angle; x2=[ω z ω y ω x ] represents the attitude angular velocity; u=[δ z δ y δ x ] represents the rudder deflection angle; d i , i=1,2 represents the aircraft modeling error and external interference; L, D, N represent lift, drag, and lateral force respectively; V, m, g represent the flight speed, hypersonic aircraft mass, and gravity acceleration respectively; I x ,I y ,I z represents the three-axis moment of inertia; θ represents the track angle; q represents the dynamic pressure; They are respectively expressed as the pitch moment coefficient related to the pitch rudder deflection angle, the yaw moment coefficient term related to the yaw rudder deflection angle, and the roll moment coefficient term related to the roll rudder deflection angle; S0 is the reference area; α is the vector composed of the angle of attack α=[1 α α 3 ] Τ , β mx =β; is the pitching moment coefficient term related to α; is the yaw moment coefficient term related to the sideslip angle; is the rolling moment coefficient term related to the roll angle.

[0099] Step 104 , based on the attitude system model, a switching dynamic event triggering strategy is designed, and an internal dynamic variable whose convergence characteristic matches the sliding mode variable is established in the switching dynamic event triggering strategy.

[0100] Specifically, the switching dynamic event triggering strategy is expressed as:

[0101]

[0102]

[0103] In the formula, u j (t) represents the jth component of the control quantity u(t); υ j Indicates the event trigger controller that will be designed; and Indicates the start and end time of the time interval of the jth channel under the event trigger mechanism; λ i ,i=1,2,3,δ are all positive coefficients;η j Represents an internal dynamic variable; z u,j represents the difference between the event-triggered controller and the ideal control input; δ represents the constant coefficient used to switch the trigger condition.

[0104] More specifically, the expressions of other coefficients are:

[0105] The parameters used to indicate whether a dynamic event is triggered are:

[0106]

[0107] The corresponding matrix form is:

[0108] b = diag(b1 b2 b3);

[0109] The vector form of the event trigger controller is:

[0110] υ=[υ1,υ2,υ3] Τ ;

[0111] The vector form of dynamic internal variables is:

[0112] η=[η1,η2,η3] Τ ;

[0113] A variable used to express the relationship between the ideal control input and the event-triggered controller is:

[0114]

[0115] Another variable used to express the relationship between the ideal control input and the event-triggered controller is:

[0116]

[0117] The internal dynamic variable expression is:

[0118]

[0119] In the formula, c r , r=1,2,3,4 are coefficients and satisfy and represents a nonlinear function with independent variable y, which has been predefined; y = [s Τ η Τ b] Τ represents the vector required in the design process; s represents the sliding surface to be designed; η is the vector form of the internal dynamic variable; Indicates a variable whose absolute value is not greater than 1; I 3×1 represents a unit vector with 3 rows and 1 column; K1 and K2 represent the elements respectively. and A diagonal matrix; i = 1, 2, 3, j = 1, 2, 3

[0120] It is worth noting that the internal dynamic variables are bounded and satisfy η j >0, the specific analysis process is as follows:

[0121] When researching | i (t)|<δ i In case of:

[0122]

[0123] By establishing the auxiliary variable η′, we can obtain:

[0124]

[0125] Assume that the auxiliary variable η′ reaches η′<0 for the first time, at which time t=t′. Considering has the same sign as η′. By studying the changing trend of η′(t0), we know that its derivative will reach This means that η′ will not continue to decrease because it is already very close to 0. Moreover, this will lead to Further explanation This contradicts the above assumption. Therefore, we can get η>η′≥0.

[0126] Since η>0, the following inequality can be established:

[0127]

[0128] Taking all the cases into consideration, η>0 holds for all u(t) cases.

[0129] Therefore, we can get

[0130]

[0131] Step 106, constructing a fixed-time sliding mode controller by adopting a fixed-time control method according to the angle tracking error and the reaching law; constructing an event-triggered fixed-time sliding mode controller according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller.

[0132] Specifically, define the angle command x c =[x 1c x 2c x 3c ] Τ , then the angle tracking error is expressed as:

[0133] z=x1-x c =[z1 z2 z3] Τ ;

[0134] According to the angle tracking error z and the derivative function of the error variable Design sliding surface s = [s1 s2 s3] Τ , the expression is:

[0135]

[0136] The derivative function of the sliding surface is obtained as:

[0137]

[0138] The sliding surface is made to converge within a fixed time through the reaching law. The reaching law can be designed as:

[0139]

[0140] Then, the fixed-time sliding mode controller function is designed as:

[0141]

[0142] In the formula, k r , r = 2, 3, 4 and μ j , j = 1, 2, 3, 4 are coefficients, satisfying and

[0143] According to the switching dynamic event triggering strategy in step 104, it can be obtained that:

[0144]

[0145] In the formula: when |u j |>δ j,1 =0 holds true.

[0146] Obviously, we can get:

[0147] u=K1υ-K2η;

[0148] The event-triggered fixed-time sliding mode controller is expressed as:

[0149]

[0150] Where υ0(t) represents the control input; x c =[x 1c x 2c x 3c ] Τ Indicates angle command; s = [s1 s2 s3] Τ represents the sliding surface; k and ε are coefficients; Indicates a positive designable coefficient.

[0151] The attitude control input is expressed as:

[0152]

[0153] Where: Ψ(y,s; k3,k4) represents a function with independent variables y and s, which has been defined in advance; k0 and k1 are coefficients, k1>L1+D2, k3=c3, k4=c4. L1 and D2 represent the upper bounds of the absolute value of the interference and the reciprocal of the interference respectively.

[0154] It is worth mentioning that in the actual application of the designed controller, the sign function is replaced by the hyperbolic tangent function. This can suppress the chattering problem and prevent the Zeno phenomenon to a certain extent.

[0155] Step 108, obtaining error information of the aircraft in real time, and using the error information as an input event to trigger a fixed-time sliding mode controller to perform calculations to obtain attitude control input.

[0156] Step 110, adjusting and controlling the attitude angle of the aircraft through attitude control input.

[0157] The above-mentioned event-triggered fixed-time control method and system for a hypersonic aircraft constructs a control-oriented attitude system model for a hypersonic aircraft; then, based on the attitude system model, a switching dynamic event triggering strategy is designed, and internal dynamic variables whose convergence characteristics match the sliding mode variables are established in the switching dynamic event triggering strategy; a fixed-time sliding mode controller is constructed using a fixed-time control method according to the angle tracking error and the convergence law; an event-triggered fixed-time sliding mode controller is constructed according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller; the error information of the aircraft is obtained in real time, and the error information is input into the event-triggered fixed-time sliding mode controller for calculation to obtain the attitude control input; and the attitude angle of the aircraft is adjusted and controlled through the attitude control input.

[0158] The present application establishes a multivariable sliding surface by using the angle tracking error and the derivative function of the error variable; the sliding surface can be converged within a fixed time by designing the convergence law; based on the designed event-triggered fixed-time control method, the controller can be directly obtained without complex hierarchical design. Based on the control-oriented attitude system model, an ideal fixed-time sliding mode controller can be designed to ensure that the aircraft angle tracking error converges within a fixed time, and the time upper bound is independent of the system initial value; the event trigger mechanism adopts switching design and dynamic internal dynamic variable design, which can obtain a better trigger strategy, reduce the control quantity update frequency under the premise of ensuring control performance, and save aircraft system resources. The complete hypersonic aircraft event-triggered fixed-time sliding mode controller consists of an ideal fixed-time sliding mode controller and a switching dynamic event triggering strategy. Based on the aircraft error information obtained in real time, the attitude control input can be directly obtained. After the control input is input to the hypersonic aircraft, the aircraft attitude can track the set angle command within a fixed time, ensuring that the aircraft attitude is fast and stable and saving system resources by reducing the controller update frequency. The present invention uses fixed-time control technology and event-triggered strategy to achieve rapid attitude control and stabilization of hypersonic deformable aircraft, ensure rapid tracking of the aircraft attitude and effective conservation of aircraft system resources, solve the difficult problem of designing a strong and robust control system with low resource consumption for hypersonic deformable aircraft, and has faster tracking speed, higher tracking accuracy and lower resource consumption.

[0159] It should be understood that although Figure 1 The steps in the flowchart are shown in sequence as indicated by the arrows, but these steps are not necessarily executed in the order indicated by the arrows. Unless otherwise specified in this document, there is no strict order restriction for the execution of these steps, and these steps can be executed in other orders. Moreover, Figure 1 At least part of the steps may include multiple sub-steps or multiple stages. These sub-steps or stages are not necessarily executed at the same time, but can be executed at different times. The execution order of these sub-steps or stages is not necessarily sequential, but can be executed in turn or alternately with other steps or at least part of the sub-steps or stages of other steps.

[0160] In one of the embodiments, two simulation calculations are provided to verify and illustrate the effectiveness and superiority of the present invention.

[0161] The first simulation example verifies the effectiveness of the designed event-triggered finite time control method, and the second simulation example compares other methods to illustrate the superiority of the present method. The following is a detailed description of the embodiments to which the present method is applied.

[0162] For all simulation cases, the controller coefficients are chosen as follows:

[0163] The coefficient settings for the sliding surface are shown in Table 1:

[0164] Table 1 Sliding surface coefficients and values

[0165]

[0166] The coefficient settings for the controller are shown in Table 2:

[0167] Table 2 Controller coefficients and values

[0168]

[0169] The coefficient settings for the event trigger strategy are shown in Table 3:

[0170] Table 3 Event trigger strategy coefficients and values

[0171]

[0172] The initial value settings for hypersonic flight are shown in Table 4:

[0173] Table 4 Hypersonic flight values

[0174]

[0175] (1) Simulation example settings

[0176] The present invention sets up a calculation example 1 to verify the effectiveness of the designed event-triggered fixed-time control method. Therefore, a square-wave angle command is used to input the controller calculation program to obtain the control input. At the same time, according to the set flight environment and initial value, a hypersonic flight simulation is performed to obtain a real-time state response. The curve difference between the state response and the angle command is observed and compared to verify the effectiveness of the method designed by the present invention and the rationality of this embodiment.

[0177] In this example, the aircraft angle command is set as:

[0178]

[0179] β c =0°,0<t<10;

[0180]

[0181] The instructions are generated through the following second-order links:

[0182]

[0183] The purpose of simulation example 2 is to compare this method with other commonly used event triggering methods to verify the advantages of this method in terms of control performance and reducing resource overhead. The other methods are fixed threshold event triggering (FTETM), relative threshold event triggering (RTETM) and switching event triggering (SETM). The simulation example will verify the tracking accuracy and the reduction of controller update times to illustrate the superiority of this method.

[0184] In this embodiment, the aircraft angle command is set as:

[0185] α c =10°;

[0186] β c =0°;

[0187] γ Vc =12°;

[0188] The second-order servo system used to generate commands is the same as that of Simulation Example 1.

[0189] (2) Analysis of simulation results

[0190] 2. Results Analysis

[0191] The simulation results of simulation examples 1 and 2 are shown in Figure 3-Figure 11 As shown in Table 5-6. In the figure, Command represents the attitude angle tracking command, that is, the command of the angle of attack, sideslip angle and roll angle; Angles represents the actual attitude angle response; times(s) represents the flight simulation time and its unit is seconds; deg represents the angle unit is degree.

[0192] The results of simulation example 1 are as follows Figures 3 to 6 As shown in the figure. Through observation and analysis, it is found that the angular response of the aircraft achieves ideal tracking of the reference command. At the same time, the aircraft undergoes continuous deformation. In this hypersonic flight phase, the proposed event-triggered fixed-time sliding mode controller can quickly stabilize the aircraft's attitude and effectively reduce the controller update frequency. Figure 3 The variation curves of the three-channel attitude angles are shown. The simulation results show that all attitude angles can reach the ideal angle within 2; Figure 4 The angular velocity variation curve corresponding to the attitude angle is shown; Figure 5 The rudder angle variation curve and the continuous 2-time deformation process are shown; Figure 6 The designed event trigger mechanism is used to demonstrate the controller update time and the time interval between two controller update times. The simulation results show that the controller update frequency is significantly reduced, and the pressure of data transmission inside the flight control system is effectively alleviated, thereby extending the life of the servo system.

[0193] The simulation results of simulation example 2 are as follows: Figure 7-11As shown in Table 5-6. Figure 7 , Figure 8 As shown in Table 5, the hypersonic flight simulation verifies that the method designed by the present invention (switching dynamic event triggering strategy, SDETM) achieves the best balance in flight control performance and resource consumption compared with traditional event triggering methods, such as fixed threshold event triggering (FTETM), relative threshold event triggering (RTETM) and switching event triggering (SETM). First, by comparing with the continuous time flight simulation, it can be found that the angle response error is only slightly reduced after using the designed event triggering mechanism. Taking the angle of attack as an example, it is reduced by 0.004°, but the number of controller adjustments is only 33.9% of that in continuous time, which shows that the control performance does not show obvious performance loss, but the resource overhead of flight control is effectively reduced. The statistics in the table are the average tracking error within the final 3 seconds. By carefully analyzing Tables 5 and 6, it can be found that the relative threshold triggering has obvious performance loss in the entire tracking process, but it also brings the least number of controller updates. Fixed threshold event triggering has the most controller updates and the highest control accuracy. At the same time, combined with the continuous time condition and the four trigger mechanisms, it is shown that the interval between two time triggers does affect the accuracy of angle tracking and the rapidity of tracking error convergence, but by reasonably designing the trigger mechanism and adjusting the event trigger conditions, time and interval between two triggers, the impact on control performance can be effectively controlled, and the advantages of the event trigger mechanism can be fully utilized. Combining the simulation data of the four methods, it can be found that the designed switching dynamic event trigger has the most ideal performance balance.

[0194] Table 5 Tracking error comparison

[0195]

[0196] Table 6 Comparison of controller update times

[0197]

[0198] In one embodiment, Fig.12 As shown, a hypersonic aircraft fixed-time control system based on event triggering is provided, including: an attitude system model building module 202, a switching dynamic event triggering strategy design module 204, a fixed-time sliding mode controller building module 206, an event-triggered fixed-time sliding mode controller building module 208, an attitude control input calculation module 210 and an attitude adjustment control module 212, wherein:

[0199] The attitude system model building module 202 is used to build an attitude system model for control of a hypersonic vehicle;

[0200] A switching dynamic event trigger strategy design module 204 is used to design a switching dynamic event trigger strategy based on the posture system model, and establish an internal dynamic variable whose convergence characteristics match the sliding mode variable in the switching dynamic event trigger strategy;

[0201] A fixed time sliding mode controller construction module 206 is used to construct a fixed time sliding mode controller using a fixed time control method according to an angle tracking error and a reaching law;

[0202] An event-triggered fixed-time sliding mode controller construction module 208 is used to construct an event-triggered fixed-time sliding mode controller according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller;

[0203] The attitude control input calculation module 210 is used to obtain the error information of the aircraft in real time, and trigger the fixed-time sliding mode controller to calculate the error information input event to obtain the attitude control input;

[0204] The attitude adjustment control module 212 is used to adjust and control the attitude angle of the aircraft through attitude control input.

[0205] For the specific definition of the event-triggered hypersonic aircraft fixed time control system, please refer to the definition of the event-triggered hypersonic aircraft fixed time control method above, which will not be repeated here. Each module in the above-mentioned event-triggered hypersonic aircraft fixed time control system can be implemented in whole or in part by software, hardware and their combination. The above-mentioned modules can be embedded in or independent of the processor in the computer device in the form of hardware, or can be stored in the memory in the computer device in the form of software, so that the processor can call and execute the operations corresponding to the above modules.

[0206] Those skilled in the art can understand that all or part of the processes in the above-mentioned embodiment methods can be completed by instructing the relevant hardware through a computer program, and the computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to memory, storage, database or other media used in the embodiments provided in this application can include non-volatile and / or volatile memory. Non-volatile memory can include read-only memory (ROM), programmable ROM (PROM), electrically programmable ROM (EPROM), electrically erasable programmable ROM (EEPROM) or flash memory. Volatile memory can include random access memory (RAM) or external cache memory. As an illustration and not limitation, RAM is available in many forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), double data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), synchronous link (Synchlink) DRAM (SLDRAM), memory bus (Rambus) direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM).

[0207] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0208] The above-described embodiments only express several implementation methods of the present application, and the descriptions thereof are relatively specific and detailed, but they cannot be construed as limiting the scope of the present invention. It should be noted that, for a person of ordinary skill in the art, several modifications and improvements may be made without departing from the concept of the present application, and these all belong to the protection scope of the present application. Therefore, the protection scope of the present application shall be subject to the attached claims.

Claims

1. A method for controlling a hypersonic vehicle at a fixed time based on event triggering, characterized in that: The method comprises: Construct a control-oriented attitude system model for hypersonic vehicles; Based on the posture system model, a switching dynamic event triggering strategy is designed, and an internal dynamic variable whose convergence characteristics match the sliding mode variable is established in the switching dynamic event triggering strategy; According to the angle tracking error and reaching law, a fixed-time sliding mode controller is constructed by using the fixed-time control method. Constructing an event-triggered fixed-time sliding mode controller according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller; Acquire error information of the aircraft in real time, input the error information into the event-triggered fixed-time sliding mode controller for calculation, and obtain attitude control input; The attitude angle of the aircraft is adjusted and controlled through the attitude control input.

2. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 1, characterized in that: Construct a control-oriented attitude system model for a hypersonic vehicle, including: Based on the original nonlinear model of the hypersonic vehicle, a control-oriented attitude system model of the hypersonic vehicle is constructed.

3. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 2, characterized in that: The original nonlinear model of the hypersonic vehicle is expressed by a nonlinear function, and the nonlinear function expression is: First, define the scalar σ, which is expressed as: Construct a nonlinear function expression: Where, σ=[σ1 σ2 … σ n ] Τ Represents the element as σ j ,j=1,...,n vector; l1,l2 represent the designable positive parameters respectively; and Ψ(σ1,σ2; l1,l2) denote pre-designed functions; ||·|| denotes the 2-norm.

4. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 3, characterized in that: The control-oriented attitude system model of the hypersonic vehicle is expressed as: in, In the formula, x1=[α β γ V ] represents the attitude angle; x2=[ω z ω y ω x ] represents the attitude angular velocity; u=[δ z δ y δ x ] represents the rudder deflection angle; d i , i=1,2 represents the aircraft modeling error and external interference; L, D, N represent lift, drag, and lateral force respectively; V, m, g represent the flight speed, hypersonic aircraft mass, and gravity acceleration respectively; I x ,I y ,I z represents the three-axis moment of inertia; θ represents the track angle; q represents the dynamic pressure; They are respectively expressed as the pitch moment coefficient related to the pitch rudder deflection angle, the yaw moment coefficient term related to the yaw rudder deflection angle, and the roll moment coefficient term related to the roll rudder deflection angle; S0 is the reference area; α is the vector composed of the angle of attack α=[1 α α 3 ] Τ , β mx =β; is the pitching moment coefficient term related to α; is the yaw moment coefficient term related to the sideslip angle; is the rolling moment coefficient term related to the roll angle.

5. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to any one of claims 1 to 4, characterized in that: The switching dynamic event triggering strategy is expressed as: Where uj(t) represents the jth component of the control variable u(t); j Indicates the event trigger controller that will be designed; and Indicates the start and end time of the time interval of the jth channel under the event trigger mechanism; λ i ,i=1,2,3,δ are coefficients; η j Represents an internal dynamic variable; z u,j represents the difference between the event-triggered controller and the ideal control input; δ represents a constant coefficient for switching the trigger condition.

6. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 5, characterized in that: The internal dynamic variable expression is: In the formula, c r , r=1,2,3,4 are coefficients; Represents a nonlinear function with independent variable y; y = [s Τ η Τ b] Τ represents the vector required in the design process; s represents the sliding surface to be designed; η is the vector form of the internal dynamic variable.

7. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 6, characterized in that: According to the angle tracking error and reaching law, a fixed-time sliding mode controller is constructed using a fixed-time control method, including: The angular tracking error is expressed as: z=x1-x c [z1 z2 z3] Τ 4 According to the angle tracking error z and the derivative function of the error variable Design sliding surface s = [s1 s2 s3] Τ , the expression is: The sliding surface is made to converge within a fixed time through the reaching law. The reaching law expression is: Then, the fixed-time sliding mode controller function is designed as: In the formula, k r ,r=2,3,4 and μ j ,j=1,2,3,4 are coefficients.

8. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 7, characterized in that: The event-triggered fixed-time sliding mode controller is expressed as: Where υ0(t) represents the control input; x c =[x 1c x 2c x 3c ] Τ Indicates angle command; s = [s1 s2 s3] Τ represents the sliding surface; k and ε are coefficients; Indicates a positive designable coefficient.

9. The method for controlling a hypersonic vehicle at a fixed time based on event triggering according to claim 8, characterized in that: The posture control input is expressed as: Where: Ψ(y,s; k3,k4) represents the function with independent variables y and s; k0 and k1 are coefficients, k1>L1+D2, k3=c3, k4=c4; L1 and D2 represent the upper bounds of the absolute values ​​of the interference and the reciprocal of the interference respectively.

10. A hypersonic vehicle fixed time control system based on event triggering, characterized in that: The system comprises: The attitude system model building module is used to build the attitude system model for control of hypersonic aircraft; A switching dynamic event trigger strategy design module is used to design a switching dynamic event trigger strategy based on the posture system model, and establish an internal dynamic variable whose convergence characteristics match the sliding mode variable in the switching dynamic event trigger strategy; A fixed-time sliding mode controller building module is used to build a fixed-time sliding mode controller using a fixed-time control method according to an angle tracking error and a reaching law; An event-triggered fixed-time sliding mode controller construction module is used to construct an event-triggered fixed-time sliding mode controller according to the switching dynamic event triggering strategy and the fixed-time sliding mode controller; An attitude control input calculation module is used to obtain error information of the aircraft in real time, input the error information into the event-triggered fixed-time sliding mode controller for calculation, and obtain attitude control input; The attitude adjustment control module is used to adjust and control the attitude angle of the aircraft through the attitude control input.

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