A method for controlling the attitude of a hypersonic vehicle during a reentry phase at a specified time

Abstract

The application discloses a specified time attitude control method for a hypersonic vehicle in a reentry phase, and comprises the following steps: according to the atmospheric environment and the self tilting steering maneuvering characteristics of the hypersonic vehicle in the reentry phase, an attitude angle dynamics model facing control is established, and a second-order error dynamics model is derived from the attitude angle dynamics model; considering the possible actuator faults and external disturbances in the flight process, the actuator faults and the external disturbances are injected into the error dynamics; based on the attitude angle dynamics model facing control, a specified time control item is designed according to the procedure of the backstepping method, including two steps: in the first step, a virtual control law is designed for the attitude angle subsystem, and in the second step, a specified time control item is designed for an intermediate error variable; and based on a nonlinear dynamic inverse control form, a final control law is constituted. The controller in the application reduces the design complexity and improves the applicability of the attitude control system.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of hypersonic vehicle control system design, and particularly relates to a specified time attitude control method for a hypersonic vehicle in reentry phase. BACKGROUND

[0002] Currently, the sliding mode variable structure control law design with robustness and finite time convergence as advantages is relatively simple, and the design method is mature, and as a passive fault-tolerant strategy, it has been widely applied. However, the convergence time of the traditional sliding mode control method is unbounded with respect to the initial state. With the more extensive application of hypersonic vehicles in military and civilian fields, the convergence time requirement for attitude tracking is further improved. In the process of high-speed flight, if the attitude tracking cannot be completed quickly, it may lead to the failure of the flight mission. Slow convergence will cause the aircraft to be in an unstable state for a long time, increasing the risk of suffering external disturbances and internal faults. These application scenarios have put forward new requirements for the control law design, that is, a control strategy with more controllable convergence time needs to be designed.

[0003] In recent years, fixed-time controller and specified time controller design have become a research hotspot at home and abroad, and the related methods include but are not limited to: quantized fixed-time attitude controller, specified time and specified boundary controller, and specified time controller based on scale transformation function, etc. However, in the existing methods, the explicit expression of the convergence time of the fixed-time controller with respect to the control parameters is not easy to obtain or the scaling boundary is too large, and the specified time controller based on scale transformation often gives unbounded control signal when approaching the specified time, which brings certain conservatism and limitation to the control law design and engineering application. Therefore, the use of specified time stability lemma to design fractional power specified time controller has attracted more and more attention, and by using the special form of Lyapunov function, the attitude control system error can be converged to the specified size of the origin neighborhood within the specified time.

[0004] Obviously, it is very necessary to design a specified time attitude control method for a hypersonic vehicle in reentry phase, which overcomes the existing conservatism, simplifies the design method, and improves the applicability of the attitude control system. SUMMARY

[0005] The technical problem solved by the present application is: with the rapidity index of the hypersonic vehicle flight task increasing, in order to improve the robustness of the vehicle attitude control system, the vehicle attitude needs to converge to the desired trajectory as soon as possible and the convergence time is adjustable, that is, the vehicle attitude needs to be stabilized within a specified time. The present application designs a specified time attitude control method for the reentry section of a hypersonic vehicle, follows the design process of the backstepping method and the form of the nonlinear dynamic inverse controller, and realizes the specified time stabilization of the attitude error. Unlike existing methods, the power form controller used in the present application overcomes the conservatism of existing methods, improves the applicability of the attitude control system, simplifies the design method, and avoids singularity.

[0006] The specified time attitude control method for the reentry section of a hypersonic vehicle of the present application comprises the following steps:

[0007] S1: According to the atmospheric environment of the reentry section of a hypersonic vehicle and the characteristics of the tilt steering maneuver of the vehicle itself, a control-oriented attitude angle dynamics model is established, and a second-order error dynamics model is derived therefrom;

[0008] S2: Considering the possible occurrence of equal actuator faults and external disturbances during flight, they are injected into the error dynamics;

[0009] S3: Based on the error dynamics model established in step S2, the specified time control term is designed following the process of the backstepping method, including two steps: first, a virtual control law is designed for the attitude angle subsystem , and second, a specified time control term is designed for the intermediate error variable ;

[0010] S4: Based on the form of the nonlinear dynamic inverse controller, the final control law is constructed .

[0011] The beneficial effects of the present application are:

[0012] (1) The attitude control method based on the specified time control theory designed in the present application adopts the backstepping method and combines the structure of the nonlinear dynamic inverse controller, which can directly obtain a controller that satisfies the specified time stabilization lemma, reducing the design complexity.

[0013] (2) The fractional power form controller used in the present application can directly adjust the upper bound of the convergence time through the adjustment of a single parameter, and can avoid the singularity problem by adding a quadratic term in the controller instead of using a filter or a piecewise continuous function, having the characteristics of simple design method and strong applicability. BRIEF DESCRIPTION OF DRAWINGS

[0014] In order to more clearly illustrate the technical solutions in the embodiments of the present application or in the prior art, the following will briefly introduce the drawings needed in the embodiments. Obviously, the drawings described below only illustrate some of the embodiments of the present application, and for those skilled in the art, other drawings can also be obtained from these drawings without creative effort.

[0015] Figure 1 Time attitude control method flow chart for reentry phase of hypersonic vehicle of the present application;

[0016] Figure 2 Attitude control effect simulation schematic diagram of the controller of the present application under normal scenario, wherein (a) shows the attitude trajectory of the vehicle under the action of the specified time controller, and (b) shows the tracking error curve of the vehicle under the action of the specified time controller;

[0017] Figure 3 Attitude control effect simulation schematic diagram of the controller of the present application under external disturbance, wherein (a) shows the attitude trajectory under external disturbance, and (b) shows the attitude error under external disturbance;

[0018] Figure 4 Control effect comparison simulation schematic diagram of the controller of the present application under different parameter settings when there is external disturbance, wherein (a) shows the attitude tracking curve of the angle of attack under different parameters, (b) shows the attitude tracking curve of the sideslip angle under different parameters, (c) shows the attitude tracking curve of the roll angle under different parameters, (d) shows the error curve of the angle of attack under different parameters, (e) shows the error curve of the sideslip angle under different parameters, and (f) shows the error curve of the roll angle under different parameters;

[0019] Figure 5 Control effect simulation schematic diagram of the controller of the present application under external disturbance and actuator failure, wherein (a) shows the attitude trajectory under external disturbance and actuator failure, and (b) shows the attitude error under external disturbance and actuator failure;

[0020] Figure 6 Rudder deflection simulation schematic diagram of the controller of the present application under external disturbance and actuator failure, wherein (a) shows the control command and actual rudder deflection of the left aileron, (b) shows the control command and actual rudder deflection of the elevator, (c) shows the control command and actual rudder deflection of the right aileron, and (d) shows the control command and actual rudder deflection of the lower elevator. DETAILED DESCRIPTION

[0021] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only some of the embodiments of the present application, but not all the embodiments.

[0022] As Figure 1 shown, the method for hypersonic vehicle reentry segment specified time attitude control comprises the following steps: first, considering the atmospheric environment of the reentry segment of the vehicle, an aerodynamic model of the reentry segment attitude control of the vehicle is established and an error dynamics model is derived; second, the actuator faults and external disturbances that may be encountered during flight are injected into the above dynamics model; then, based on the specified time control theory and the backstepping method design steps, the controller design process is divided into two stages, the first stage is to design a virtual controller , which ensures the specified time convergence of the attitude angle subsystem, i.e. the attitude error tends to zero, and the second stage is to design a specified time control term for the intermediate error variable; finally, based on the nonlinear dynamic inverse control method, the complete controller form is given. As can be seen from the above steps, the controller in the present application is designed by backstepping method, based on the specified time control theory, combined with the structure of the nonlinear dynamic inverse controller, the controller satisfying the specified time stability lemma can be directly obtained, which reduces the design complexity and improves the applicability of the attitude control system.

[0023] The method for hypersonic vehicle reentry segment specified time attitude control comprises the following steps:

[0024] S1: According to the atmospheric environment of the reentry segment of the hypersonic vehicle and the tilt steering maneuver characteristics of the vehicle, a control-oriented attitude angle dynamics model is established, and a second-order error dynamics model is derived therefrom, wherein the control-oriented attitude angle dynamics model is:

[0025] (1)

[0026] wherein is an attitude angle vector, wherein , and represent the angle of attack, the sideslip angle and the roll angle respectively; represent the roll, yaw and pitch angular rates of the vehicle rotating around the three axes of the body coordinate system respectively; is the control moment of the vehicle, wherein , , represent the roll, pitch and yaw moments respectively, is the attitude angle rate, is the roll, yaw and pitch angular acceleration vector. The matrix , , is represented as:

[0027] ,

[0028] ,

[0029] ,

[0030] where, , = Ixx Iyy Izz , , represent the moments of inertia.

[0031] The process of obtaining the second order error dynamics model includes:

[0032] In actual cases, the mapping relationship between the deflection of the control surface and the control moment of the aircraft can be expressed as where, , , is the actual deflection of the control surface, represents the number of control surfaces. The elements of the control allocation matrix satisfy ; where, , and are the deflection amounts of the aileron, elevator and rudder, respectively. Among them,

[0033] ,

[0034] where, denotes the dynamic pressure, characterizes the air density, is the flight speed. and represent the reference length and reference area of the aircraft. The moment coefficients in the roll, pitch and yaw directions are expressed as , and , and the partial derivatives of each moment coefficient with respect to the deflection of the three rudders are denoted as ( , ).

[0035] Let the reference trajectory be represented by , then the attitude tracking error is denoted as . Combined with the aforementioned control-oriented attitude angle dynamics model, the error dynamics model

[0036] (2)

[0037] where, is the derivative of the attitude tracking error, is the derivative of the attitude angular rate tracking error, ,​ is the derivative of the desired trajectory, , , , is the second derivative of the desired trajectory.

[0038] S2: Considering the actuator failure, external disturbance and other factors in the flight process, inject into the second-order error dynamics model. Among them, the actuator failure can be rudder failure, stuck and other problems. Including:

[0039] Rewrite the second-order error dynamics model formed in step S1 as:

[0040] (3)

[0041] where, represents the lumped external disturbance vector, is the control matrix considering failure, expressed as:

[0042] (4)

[0043] (5)

[0044] (6)

[0045] where, and represent unmodeled dynamics, represent external disturbances, represent the system matrix under failure, represent the control matrix under failure, represent the trigger function. When the actuator failure occurs, and mutate into and . Assuming that the failure triggers at time, this process can be represented by the trigger function , is expressed as:

[0046] (7)

[0047] S3: Based on the attitude angle dynamics model established in step S1, the time control item is designed according to the process of backstepping method. The design process is divided into two steps. In the first step, the virtual control law of the attitude angle subsystem is designed so that the attitude angle converges within a specified time, and then the specified time control item of the intermediate error variable is designed so that the entire second-order system has specified time stability. The detailed implementation steps are as follows:

[0048] virtual control law The design procedure is:

[0049] For a given vector and a constant , define the operation:

[0050] ,

[0051] where represents the sign function.

[0052] virtual control law is designed as:

[0053] (8)

[0054] (9)

[0055] where , , , are controller parameters, , , , is a specified convergence time.

[0056] Define the Lyapunov function:

[0057] (10)

[0058] Taking the derivative, we have:

[0059] (11)

[0060] Define the intermediate error variable as It can be proved that:

[0061] (12)

[0062] This shows that has a specified time stability.

[0063] Second step, the specified time control term is designed as:

[0064] (13)

[0065] (14)

[0066] where , , is the controller parameter, satisfying Similarly, the synchronization definition can be:

[0067] (15)

[0068] S4: the control-oriented attitude angular dynamics model obtained by combining S1 and S2, using the specified time control term designed in step S3, giving the final control law based on the nonlinear dynamic inverse control form, the detailed implementation steps are as follows:

[0069] Let When the system has no external disturbance, the equivalent control term can be obtained:

[0070] (16)

[0071] Where the generalized inverse matrix satisfies , is a three-order unit matrix.

[0072] According to the nonlinear dynamic inverse control form, the final control law is designed as:

[0073] (17)

[0074] The derivative is:

[0075] (18)

[0076] It can be proved that:

[0077] (19)

[0078] This shows that the entire system has specified time stability.

[0079] It should be understood that all variables in the present application have a dot above them, which is the derivative of the variable, unless the derivative of the variable has an actual physical meaning.

[0080] Next, taking the attitude tracking of a certain hypersonic vehicle as an example, the effectiveness of the method proposed in the present application is illustrated. Among them, the ideal attitude trajectory varies in the form of sine, constant, and quadratic function in three channels respectively. The mass of the vehicle itself is 1200 kg, the reference area is 334.7㎡, the reference length is 24.384 m, the moment of inertia = 434270 , =961220 , =1131541 . The four actuators are defined as , where the elements represent the left aileron, up rudder, right aileron, and down rudder respectively. satisfy , set to [0.25, 0.25, -0.25, -0.25; 0, 0.5, 0, 0.5; 0.5, 0, 0.5, 0]. The servo limit is [-20°, 20°]. The relevant parameters of the initial flight conditions are shown in Table 1 below:

[0081] Table 1 Initial flight conditions

[0082]

[0083] The posture tracking effect in normal scenes is as follows Figure 2 As shown, Figure 2 (a) shows the attitude trajectory of the aircraft under the action of the controller at a specified time, and (b) shows the tracking error curve of the aircraft under the action of the controller at a specified time. The controller parameters are selected as follows: s, , , , , .from Figure 2 It can be seen that this method can make the tracking error converge to the origin neighborhood within the specified convergence time, effectively achieving fast posture tracking. The posture tracking effect under external interference is as follows Figure 3 As shown in Figure 2, (a) shows the posture trajectory under external interference, and (b) shows the posture error under external interference. The error form is:

[0084]

[0085] in , is the time variable.

[0086] from Figure 3 It can be seen that the controller in the present invention has strong robustness to external interference. Figure 4 Shown Comparison of control effects under different controller parameters, Figure 4 (a), (b), (c), (d), (e), and (f) show the attitude tracking curves and error curves of the angle of attack, sideslip angle, and roll angle under different parameters. Figure 4 It can be seen that increasing the parameter or reduce It can reduce the tracking error under external interference, but it may cause oscillation. The posture tracking effect under external interference and actuator failure is as follows: Figure 5The figures are shown, wherein (a) shows the attitude trajectory under external disturbance and actuator failure, (b) shows the attitude error under external disturbance and actuator failure. The actuator failure includes: a fixed deviation of the left aileron at 5 s, the upper rudder is stuck at -2 deg at 10 s, and the right aileron is 60% invalid at 15 s. From Figure 5 It can be seen that the controller in the application has strong robustness to external disturbance and actuator failure, and has passive fault-tolerant properties. Figure 6 (a), (b), (c) and (d) of the figures show the control command and actual rudder deflection of the left aileron, upper rudder, right aileron and lower rudder. From Figure 6 It can be seen that at the moment of each failure, the remaining healthy actuators will immediately respond to compensate, achieving passive fault-tolerant effect.

[0087] According to the above analysis and description, it can be seen that the controller proposed in the application effectively achieves the requirements of the specified time attitude control, and has passive fault-tolerant properties, and has the characteristics of simple design method, strong adaptability, high reliability and the like.

[0088] The above only describes the preferred embodiments of the application, and is not used to limit the application, and any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the application shall be included in the protection scope of the application.

Claims

1. A method for high supersonic vehicle reentry segment specified time attitude control, characterized in that, The method comprises the following steps: S1: according to the atmospheric environment of the hypersonic vehicle reentry phase flight and the tilt steering maneuver characteristics of the hypersonic vehicle itself, a control-oriented attitude angle dynamics model is established, and a second-order error dynamics model is derived therefrom; S2: considering the actuator failure and external disturbance in the flight process, the actuator failure and the external disturbance are injected into the second-order error dynamics model; S3: Based on the attitude angle dynamics model for control established in step S1, the time control item is specified in accordance with the flow design of backstepping method, including two steps: first, a virtual control law is designed for the attitude angle subsystem , second, a time control item is designed for the intermediate error variable ; S4: combining the control-oriented attitude angle dynamics model obtained in S1 and S2, using the specified time control term designed in S3, and based on the nonlinear dynamic inverse control form, to form the final control law ; The process of specifying the time control item in step S3 follows the process design of the backstepping method: For a given vector and constant , define the operation: ， wherein represents a symbol function; Virtual control law Designed to: （8） （9） wherein, , , , is a controller parameter, , , , is a specified convergence time; Define the intermediate error variable as , the time control item is designed as: （10） （11） wherein , , is a controller parameter, .

2. The hypersonic vehicle re-entry phase specified time attitude control method of claim 1, wherein, The control-oriented attitude angle dynamics model established in step S1 is: （1） in, is the attitude angle vector, where 、 and denote the angle of attack, sideslip angle, and roll angle respectively; Indicates the roll, yaw and pitch angular rates of the aircraft around the three axes of the body coordinate system; is the vehicle control torque, where , , represent the rolling, pitching and yaw moments respectively, is the attitude angular rate, are the roll, yaw and pitch accelerations, matrices , , Expressed as: ， ， ， wherein represent the moment of inertia;​​​​​ In practical cases, the mapping between the control surface deflections and the aircraft control moments is represented as where , is the actual control surface deflection, where represents the number of control surfaces, and the elements of the control allocation matrix satisfy ; where , and are the deflection amounts of the aileron, elevator and rudder, respectively, where,​ ， wherein, denotes the dynamic pressure, characterizes the air density, is the flight speed, with denote the reference length and the reference area of the aircraft, the moment coefficients in roll, pitch, yaw direction are denoted by , with , the partial derivatives of each moment coefficient with respect to the three rudder deflections are denoted by ( , ), The reference trajectory is denoted by The pose tracking error is denoted by The second-order error dynamics model is derived in combination with the control-oriented pose angular dynamics model: （2） wherein is a derivative of the attitude tracking error, is a derivative of the attitude angular velocity tracking error, , is a derivative of the desired trajectory, , , , is a second derivative of the desired trajectory.

3. The hypersonic vehicle re-entry phase specified time attitude control method of claim 2, wherein, The injection into the error dynamics in step S2 comprises: The second-order error dynamics model formed in S1 is rewritten as: （3） wherein, represents a lumped external disturbance vector, is the control matrix considering faults, denoted as: （4） （5） （6） where with representing unmodeled dynamics, representing external disturbances, representing the system matrix under fault, representing the control matrix under fault, representing the triggering function when actuator fault occurs, and jump to and respectively, assuming the fault is triggered at time instant, this process is represented by the triggering function is represented as:​ （7）。 4. The hypersonic vehicle re-entry phase specified time attitude control method of claim 3, wherein, In step S4, the final control law The structure is: （12） where the generalized inverse matrix satisfies , is a three-order identity matrix.

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