Hypersonic flight vehicle reentry period specified time attitude control method

Abstract

The invention discloses a hypersonic flight vehicle reentry section specified time attitude control method, which comprises the following steps: establishing a control-oriented attitude angle dynamic model according to the atmospheric environment and self tilt steering maneuvering characteristics of the hypersonic flight vehicle in the reentry section, and deriving a second-order error dynamic model; an actuator fault and external interference which may occur in the flight process are considered and are injected into the error dynamics; based on a control-oriented attitude angle dynamic model, a specified time control item is designed by following the process of a backstepping method, and the method comprises the following two steps: 1, designing a virtual control law for an attitude angle subsystem, and 2, designing a specified time control item for an intermediate error variable; and forming a final control law based on a nonlinear dynamic inverse control form. According to the controller, the design complexity is reduced, and the applicability of the attitude control system is improved.

Description

Technical Field

[0001] The present invention belongs to the field of hypersonic aircraft control system design, and in particular, relates to a hypersonic aircraft reentry phase designated time attitude control method. Background Art

[0002] At present, the design of sliding mode variable structure control law with the advantages of robustness and finite time convergence is relatively simple, the design method is mature, and it has been widely used as a passive fault-tolerant strategy. However, the convergence time of the traditional sliding mode control method is unbounded with respect to the initial state. With the wider application of hypersonic aircraft in military and civilian fields, the convergence time requirements for attitude tracking are further increased. During high-speed flight, if attitude tracking cannot be completed quickly, the flight mission may not be successfully executed. Slow convergence will cause the aircraft to be in an unstable state for a long time, increasing the risk of external disturbances and internal failures. These application scenarios put forward new requirements for the design of control laws, that is, it is necessary to design a control strategy with a more controllable convergence time.

[0003] In recent years, the design of fixed-time controllers and specified-time controllers has become a research hotspot at home and abroad. Related methods include but are not limited to: quantized fixed-time attitude controllers, specified-time specified-boundary controllers, and specified-time controllers based on scaling functions. However, in existing methods, the explicit expression between the convergence time of fixed-time controllers and control parameters is difficult to obtain or the scaling boundary is too large. The specified-boundary control based on scaling often gives unbounded control signals when approaching the specified time, which brings certain conservatism and limitations to the control law design and engineering applications. Therefore, the use of the specified-time stability lemma to design fractional power specified-time controllers has received more and more attention. By using the special form of the Lyapunov function, the attitude control system error can be made to converge to a specified size of the origin neighborhood within a specified time.

[0004] Obviously, it is necessary to design a method for attitude control at a specified time during the reentry phase of a hypersonic vehicle, while overcoming the existing conservatism, simplifying the design method, and improving the applicability of the attitude control system. Summary of the invention

[0005] The technical problem solved by the present invention is: as the rapidity index of the hypersonic aircraft flight mission increases, in order to improve the robustness of the aircraft attitude control system, the aircraft attitude needs to converge to the desired trajectory as quickly as possible and the convergence time can be adjusted, that is, the aircraft attitude needs to be stabilized within a specified time. The present invention designs a hypersonic aircraft reentry phase specified time attitude control method, which follows the design process of the backstepping method and the form of a nonlinear dynamic inverse controller to achieve the specified time stability of the attitude error. Different from the existing methods, the power form controller used in the present invention simplifies the design method and avoids singularity while overcoming the existing conservatism and improving the applicability of the attitude control system.

[0006] A method for controlling attitude at a specified time during reentry of a hypersonic vehicle according to the present invention comprises the following steps:

[0007] S1: According to the atmospheric environment of the hypersonic vehicle during reentry and its own tilting and turning maneuvering characteristics, a control-oriented attitude angle dynamics model is established, and a second-order error dynamics model is derived from it;

[0008] S2: Consider possible actuator failures and external disturbances during flight and inject them into the error dynamics;

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

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

[0011] Beneficial effects of the present invention:

[0012] (1) The attitude control method based on the specified time control theory designed in the present invention adopts the backstepping design method and combines it with the nonlinear dynamic inverse controller structure to directly obtain a controller that satisfies the specified time stability lemma, thereby reducing the design complexity.

[0013] (2) The present invention adopts a fractional-order power form controller. By adjusting a single parameter, the upper limit of the convergence time can be directly adjusted. The singularity problem is avoided by adding a quadratic term to the controller instead of using a filter or a piecewise continuous function. It has the characteristics of simple design method and strong applicability. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0015] Figure 1 A flow chart of a method for controlling the attitude at a designated time during the reentry phase of a hypersonic vehicle according to the present invention;

[0016] Figure 2 It is a schematic diagram of the simulation of the attitude control effect of the controller of the present invention in a normal scenario, wherein (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;

[0017] Figure 3 It is a schematic diagram of the simulation of the posture control effect of the controller of the present invention under external interference, wherein (a) shows the posture trajectory under external interference, and (b) shows the posture error under external interference;

[0018] Figure 4 Schematic diagram of simulation comparison of control effects of the controller of the present invention under different parameter settings when there is external interference, wherein (a) shows attitude tracking curves of angle of attack under different parameters, (b) shows attitude tracking curves of sideslip angle under different parameters, (c) shows attitude tracking curves of roll angle under different parameters, (d) shows error curves of angle of attack under different parameters, (e) shows error curves of sideslip angle under different parameters, and (f) shows error curves of roll angle under different parameters;

[0019] Figure 5 It is a schematic diagram of the simulation of the control effect of the controller of the present invention under external interference and actuator failure, wherein (a) shows the posture trajectory under external interference and actuator failure, and (b) shows the posture error under external interference and actuator failure;

[0020] Figure 6 The present invention is a schematic diagram of the rudder deflection simulation of the controller under external interference 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 upper rudder, (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 rudder. DETAILED DESCRIPTION

[0021] The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments.

[0022] like Figure 1 As shown in the figure, the steps of a method for attitude control of a hypersonic vehicle during the reentry phase of the present invention are as follows: first, considering the atmospheric environment of the aircraft during the reentry phase, an aerodynamic model of attitude control of the aircraft during the reentry phase is established and an error dynamic model is derived; secondly, actuator failures and external interferences that may be encountered during the flight are injected into the above dynamic 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 designs a virtual controller. , to ensure the specified time convergence of the attitude angle subsystem, that is, the attitude error tends to zero, and in the second stage, the specified time control term is designed for the intermediate error variable; finally, a complete controller form is given based on the nonlinear dynamic inverse control method. From the above steps, it can be seen that the controller in the present invention is designed using the backstepping method, based on the specified time control theory, combined with the nonlinear dynamic inverse controller structure, a controller that satisfies 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 hypersonic vehicle reentry phase designated time attitude control method of the present invention specifically comprises the following steps:

[0024] S1: According to the atmospheric environment of the hypersonic vehicle during the reentry phase and its own tilting and turning maneuvering characteristics, a control-oriented attitude angle dynamics model is established, and a second-order error dynamics model is derived from it. The control-oriented attitude angle dynamics model is:

[0025] (1)

[0026] in, is the attitude angle vector, where , and They represent the angle of attack, sideslip angle and roll angle respectively; They represent the roll, yaw and pitch angular rates of the aircraft rotating around the three axes of the body coordinate system respectively; 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 acceleration vectors. The matrix , , It is expressed as:

[0027] ,

[0028] ,

[0029] ,

[0030] in, ( , = , , ) represents the moment of inertia.

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

[0032] In actual situations, the deflection of the rudder surface and the control torque of the aircraft The mapping relationship between them can be expressed as ,in , is the actual rudder deflection, Represents the number of control surfaces. Control allocation matrix The elements satisfy ; ,in , and are the deflections of the aileron, elevator and rudder respectively.

[0033] ,

[0034] in, Indicates dynamic pressure, Characterizes the air density, For the flight speed. and Represents the reference length and reference area of ​​the aircraft. The moment coefficients in the roll, pitch, and yaw directions are expressed as , and The partial derivatives of the moment coefficients with respect to the three rudder deflections are expressed as ( , ).

[0035] The reference trajectory is The attitude tracking error is expressed as The error dynamics model can be derived by combining the aforementioned control-oriented attitude angle dynamics model

[0036] (2)

[0037] in, 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-order derivative of the desired trajectory.

[0038] S2: Consider actuator failures and external interference during flight and inject them into the second-order error dynamics model. Actuator failures can be problems such as rudder failure and jamming. Including:

[0039] The second-order error dynamics model formed in step S1 is rewritten as:

[0040] (3)

[0041] in, represents the aggregate external interference vector, is the control matrix considering faults, expressed as:

[0042] (4)

[0043] (5)

[0044] (6)

[0045] in, and represents unmodeled dynamics, represents the external disturbance, represents the system matrix under fault, represents the control matrix under fault, Represents the trigger function. When an actuator failure occurs, and Mutated to and . Assume that the fault is triggered at At this moment, the trigger function This indicates that the process It is expressed as:

[0046] (7)

[0047] S3: Based on the control-oriented attitude angle dynamics model established in step S1, design the specified time control item following the backstepping process. 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 the specified time, and then the specified time control item is designed for the intermediate error variable , so that the entire second-order system has the specified time stability. The detailed implementation steps are as follows:

[0048] Virtual Control Law The design process is:

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

[0050] ,

[0051] in, Represents a symbolic function.

[0052] Virtual Control Law Designed for:

[0053] (8)

[0054] (9)

[0055] in, , , , are controller parameters, , , , To specify the convergence time.

[0056] Define the Lyapunov function:

[0057] (10)

[0058] Taking the derivative, we have:

[0059] (11)

[0060] Define the intermediate error variable as , we can prove that:

[0061] (12)

[0062] This shows that With specified time stability.

[0063] Step 2: Specify time control items Designed for:

[0064] (13)

[0065] (14)

[0066] in, , , is a controller parameter that satisfies . Similarly, you can define synchronously:

[0067] (15)

[0068] S4: Combining the attitude angle dynamics model for control obtained in S1 and S2, using the specified time control term designed in step S3, based on the nonlinear dynamic inverse control form, the final control law is given. The detailed implementation steps are as follows:

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

[0070] (16)

[0071] The generalized inverse matrix satisfy , is the third-order identity matrix.

[0072] According to the nonlinear dynamic inverse control form, the final control law Designed for:

[0073] (17)

[0074] The derivatives are:

[0075] (18)

[0076] It can be proved that:

[0077] (19)

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

[0079] It should be understood that all variables with dots above them in the present invention are derivatives of the variables, unless the derivative of the variable has actual physical meaning.

[0080] The following example takes the attitude tracking of a hypersonic aircraft to illustrate the effectiveness of the method proposed in the present invention. The ideal attitude trajectory changes in the three channels in the form of sine, constant, and quadratic functions. The aircraft has a mass of 1200 kg, a reference area of ​​334.7 m2, a reference length of 24.384 m, and a moment of inertia of = 434270 , =961220 , =1131541 The four actuators are defined as , where the elements represent the left aileron, the upper rudder, the right aileron, and the lower rudder. 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 realizing fast posture tracking. Figure 3 As shown, (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 The tracking error under external disturbance can be reduced, but it may cause oscillation. The posture tracking effect under external disturbance and actuator failure is as follows Figure 5As shown in the figure, (a) shows the attitude trajectory under external interference and actuator failure, and (b) shows the attitude error under external interference and actuator failure. The actuator failures include: the left aileron has a fixed deviation at 5 s, the upper rudder is stuck at -2 deg at 10 s, and the right aileron fails by 60% at 15 s. Figure 5 It can be seen that the controller in the present invention is highly robust to external interference and actuator failures and has passive fault tolerance properties. Figure 6 (a), (b), (c) and (d) show the control commands and actual rudder deflection of the left aileron, up rudder, right aileron and down rudder. Figure 6 It can be seen that at the moment each fault occurs, the remaining healthy actuators will immediately respond to compensate and achieve a passive fault-tolerant effect.

[0087] According to the above analysis and description, it can be seen that the controller proposed in the present invention effectively realizes the requirements of specified time posture control, and has passive fault tolerance characteristics, and has the characteristics of simple design method, strong adaptability, and high reliability.

[0088] 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. A method for controlling the attitude of a hypersonic vehicle during a reentry phase at a specified time, characterized in that: The following steps are involved: S1: According to the atmospheric environment of the hypersonic vehicle during reentry and its own tilting and turning maneuvering characteristics, a control-oriented attitude angle dynamics model is established, and a second-order error dynamics model is derived from it; S2: Consider actuator failures and external disturbances during flight and inject them into the second-order error dynamics model; S3: Based on the control-oriented attitude angle dynamics model established in step S1, the specified time control item is designed following the backstepping process, including two steps: the first step is to design a virtual control law for the attitude angle subsystem , the second step is to design a specified time control term for the intermediate error variable ; S4: Combining the attitude angle dynamics model for control obtained in S1 and S2, using the specified time control term designed in step S3, based on the nonlinear dynamic inverse control form, the final control law is formed .

2. The hypersonic vehicle reentry phase designated time attitude control method according to claim 1, characterized in that: The control-oriented attitude angle dynamics model established in step S1 is: （1） in, is the attitude angle vector, where , and They represent the angle of attack, sideslip angle and roll angle respectively; Indicates the roll, yaw and pitch angular rates of the aircraft rotating 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 angular accelerations, matrices , , It is expressed as: ， ， ， in, ( , = , , ) represents the moment of inertia; In actual situations, the deflection of the rudder surface and the control torque of the aircraft The mapping relationship between them is expressed as ,in , is the actual rudder deflection, where Represents the number of rudders and controls the distribution matrix The elements in ; ,in , and are the deflections of the aileron, elevator and rudder respectively, where ， in, Indicates dynamic pressure, Characterizes the air density, is the flight speed, and Represents the reference length and reference area of ​​the aircraft, and the moment coefficients in the roll, pitch, and yaw directions are expressed as , and The partial derivatives of the moment coefficients with respect to the three rudder deflections are expressed as ( , ), The reference trajectory is The attitude tracking error is expressed as , combined with the control-oriented attitude angle dynamics model, the second-order error dynamics model is derived: （2） in, is the derivative of the attitude tracking error, is the derivative of the attitude angular velocity tracking error, , is the derivative of the desired trajectory, , , , is the second-order derivative of the desired trajectory.

3. The hypersonic vehicle reentry phase designated time attitude control method as claimed in claim 2, characterized in that: Injecting it into the error dynamics in step S2 includes: The second-order error dynamics model formed in S1 is rewritten as: （3） in, represents the aggregate external interference vector, is the control matrix considering faults, expressed as: （4） （5） （6） in, and represents unmodeled dynamics, represents the external disturbance, represents the system matrix under fault, represents the control matrix under fault, Represents the trigger function. When an actuator failure occurs, and Mutated to and , assuming the fault is triggered at At this moment, the trigger function This indicates that the process It is expressed as: （7）。 4. The hypersonic vehicle reentry phase designated time attitude control method as claimed in claim 3, characterized in that: Step S3 follows the process of backstepping method to designate the time control item: For a given vector And constant , define the operation: ， in, represents a symbolic function; Virtual Control Law Designed for: （8） （9） in, , , , are controller parameters, , , , To specify the convergence time; Define the intermediate error variable as , specify time control items Designed for: （10） （11） in, , , are controller parameters, .

5. The hypersonic vehicle reentry phase designated time attitude control method as claimed in claim 4, characterized in that: In step S4, the final control law The structure is: （12） Among them, the generalized inverse matrix satisfy , is the third-order identity matrix.

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