Self-adaptive anti-interference guidance and control integrated method for hypersonic flight vehicle

Abstract

The invention relates to a self-adaptive anti-interference guidance and control integrated method for a hypersonic flight vehicle, belongs to the field of flight vehicle control, and aims at the guidance and control integrated problem of the hypersonic flight vehicle with terminal angle constraint, inertial parameter time varying, parameter uncertainty and external disturbance. Establishing kinematics and dynamics models of the hypersonic aircraft with inertial parameter time varying; secondly, separating time-varying inertial parameters, and establishing a guidance and control integrated model in a strict feedback form based on kinematics and dynamical models; and finally, designing an adaptive law to estimate inertial parameters of the aircraft and boundaries of unknown interference, and based on a dynamic surface method, combining the adaptive law to design a guidance and control integrated scheme to complete the adaptive anti-interference guidance and control integrated method for the hypersonic aircraft. The method realizes comprehensive guidance control of the hypersonic flight vehicle under the constraint of the terminal angle, has the characteristic of high anti-interference capability, and is suitable for a hypersonic flight vehicle system with time-varying inertial parameters.

Description

Technical Field

[0001] The present invention belongs to the field of aircraft control, and particularly relates to an integrated method for adaptive anti-interference guidance and control of a hypersonic aircraft. Background Art

[0002] Most traditional aircraft guidance and control systems are designed based on the spectrum separation principle, separating the design of the control subsystem and the guidance subsystem. Different from general low-speed aircraft, when a hypersonic aircraft is in a hypersonic flight state, the assumption that the control loop time constant is much smaller than the guidance loop time constant is difficult to guarantee, resulting in the phenomenon that the traditional method of designing controllers for the guidance and control systems in separate loops based on the spectrum separation assumption and then making them coordinated and matched through joint debugging is prone to a decrease in terminal accuracy or even aircraft out of control. Different from the series and transfer control ideas of the controlled state of separating guidance and control first and then combining them, the concept of integrated guidance and control design is to directly generate actuator control commands according to the centroid and the state around the centroid of the aircraft and simultaneously complete the tasks of centroid guidance and attitude control around the centroid of the aircraft. The advantages of adopting integrated guidance and control design are that it no longer relies on the spectrum separation assumption, can improve system reliability; can make full use of various couplings existing in the system, thereby improving system performance as a whole; avoids the problem of repeated joint debugging in the traditional design method, and reduces the design cost. At present, there is less research on the integrated guidance and control of aircraft considering time-varying inertial parameters. However, during the flight process, the fuel consumption is significant, and there will always be changes in mass and moment of inertia, which will in turn cause certain time-variation in the entire aircraft system and may lead to system instability. Therefore, the research on the integrated guidance and control of aircraft considering time-varying inertial parameters is of great significance.

[0003] At present, the design methods for the integrated guidance and control of hypersonic vehicles mainly focus on methods such as backstepping and sliding mode. In Chinese Patent Application CN202311001454.6, a self-disturbance rejection integrated guidance and control design method is proposed. For the full-coupled six-degree-of-freedom model of the vehicle, the self-disturbance rejection control method is used to design the integrated control law, and the state observer is used to estimate and compensate the uncertain terms and external environmental disturbance terms in the model to improve the anti-disturbance ability. In Chinese Patent Application CN202311335283.0, a backstepping super-twisting sliding mode integrated guidance and control method for vehicles is proposed. Aiming at the problem of unstable guidance and control of the vehicle, a backstepping super-twisting sliding mode controller is designed, and the high-order sliding mode differentiator is used to observe the time derivative of the virtual control quantity to make the vehicle hit the target. In Chinese Patent Application CN201910432664.8, an integrated guidance and control method for vehicles based on observer technology is proposed. Combining variable structure sliding mode control and backstepping control, a multi-sliding surface backstepping controller is designed, and the second-order disturbance observer is used to observe and compensate the disturbance. This method is beneficial to prevent the actuator from saturation and improves the anti-disturbance ability and robustness of the system. The above methods have all solved the problem of mismatch between the guidance and control systems in traditional guidance and control, but have not considered the terminal angle constraint of the vehicle and the problems of mass and moment of inertia changes. In the existing integrated guidance and control design technologies considering the attack angle constraint and input saturation, for the integrated guidance and control problem of hitting a stationary target, considering the input saturation problem and attack angle constraint, a finite-time observer is used to estimate and compensate the system model uncertainty, and the terminal sliding mode control and backstepping control are used to propose a three-dimensional integrated guidance and control three-channel independent design method. The article (Wang XH, Tan CP, Cheng LP. Impact time and angle constrained integrated guidance and control with application to salvo attack[J]. Asian Journal of Control, 2020, 22: 1211-1220.) proposes a new method to estimate the impact time for the integrated guidance and control problem of hitting a stationary target. The designed control law makes the line-of-sight angle converge to the expected value within a finite time, and the missile will fly along the expected line of sight until it hits the target. The above methods consider the terminal angle constraint problem of the vehicle, but still do not consider the problem of time-varying inertial parameters.

[0004] In summary, in the scenarios of terminal angle constraint, time-varying inertial parameters, parameter uncertainty, and external disturbance, the existing methods lack an anti-disturbance integrated guidance and control scheme, and it is urgent to overcome the adaptive anti-disturbance integrated guidance and control method for hypersonic vehicles. Summary of the Invention

[0005] For the problem of integrated guidance and control of hypersonic vehicles, considering the cases of terminal angle constraints, time-varying inertial parameters, parameter uncertainties, and external disturbances, in order to overcome the deficiencies of existing technologies, the present invention provides an adaptive anti-disturbance integrated guidance and control method for hypersonic vehicles. By using a method that combines adaptive technology and dynamic surface technology, the vehicle can hit a stationary target at a desired angle, improving the attack effect of the vehicle and achieving adaptive estimation and compensation of unknown total disturbances.

[0006] To achieve the above object, the present invention adopts the following technical solutions:

[0007] An adaptive anti-disturbance integrated guidance and control method for hypersonic vehicles, comprising the following steps:

[0008] First step, establish the kinematic model and dynamic model of a hypersonic vehicle with time-varying inertial parameters;

[0009] Second step, separate the time-varying inertial parameters, and based on the kinematic model and dynamic model of the hypersonic vehicle, establish an integrated guidance and control model in strict-feedback form;

[0010] Third step, design an adaptive law to estimate the inertial parameters of the vehicle and the bounds of the unknown disturbances it suffers. Based on the dynamic surface method, combine the adaptive law to design an integrated guidance and control scheme, and complete the adaptive anti-disturbance integrated guidance and control method for hypersonic vehicles.

[0011] The beneficial effects of the present invention compared with the prior art are as follows:

[0012] Aiming at the shortcoming of the existing method lacking a solution for the time-varying inertial parameters of the vehicle, the present invention designs an integrated guidance and control scheme that combines adaptive technology and dynamic surface technology, enabling it to still hit a stationary target at a desired angle under the condition of changes in mass and moment of inertia, improving the anti-disturbance ability of the system, and being applicable to hypersonic vehicle systems with time-varying inertial parameters under multi-source disturbances. Description of the Drawings

[0013] Figure 1 Is the design flow chart of an adaptive anti-disturbance integrated guidance and control method for a hypersonic vehicle of the present invention;

[0014] Figure 2 Is the control block diagram of an adaptive anti-disturbance integrated guidance and control method for a hypersonic vehicle of the present invention;

[0015] Figure 3 Is the flight trajectory diagram of a hypersonic vehicle in a simulation experiment;

[0016] Figure 4 Is the distance schematic diagram between a hypersonic vehicle and a target in a simulation experiment;

[0017] Figure 5 It is a schematic diagram of the line-of-sight angle of a hypersonic vehicle in a simulation experiment. Among them, a is the tracking effect of the line-of-sight inclination angle, and b is the tracking effect of the line-of-sight deflection angle;

[0018] Figure 6 It is a schematic diagram of the attitude angle of a hypersonic vehicle in a simulation experiment. Among them, a is the angle-of-attack curve, b is the sideslip-angle curve, and c is the velocity-tilt-angle curve. Specific implementation mode

[0019] The present invention will be described in detail below with reference to the accompanying drawings and embodiments.

[0020] As Figure 1 , Figure 2 shown, an integrated guidance and control method for a hypersonic vehicle with adaptive anti-interference of the present invention includes the following steps:

[0021] The first step: Establish the kinematic and dynamic models of a hypersonic vehicle with time-varying inertial parameters;

[0022] The second step: Separate the time-varying inertial parameters, and establish an integrated guidance and control model in strict-feedback form based on the kinematic and dynamic models;

[0023] The third step: Design an adaptive law to estimate the inertial parameters of the vehicle and the bounds of the unknown disturbances, and based on the dynamic surface method, combine the adaptive law to design an integrated guidance and control scheme to complete the integrated guidance and control method for a hypersonic vehicle with adaptive anti-interference.

[0024] The present invention combines the kinematic and dynamic models of a hypersonic vehicle, derives an integrated guidance and control model in strict-feedback form with time-varying inertial parameters, combines adaptive and dynamic surface technologies to design an integrated guidance and control scheme, completes the integrated guidance and control method for the vehicle, and finally realizes the integrated guidance and control of a hypersonic vehicle under terminal angle constraints, with the characteristics of strong anti-interference ability, and is applicable to a hypersonic vehicle system with time-varying inertial parameters.

[0025] Specifically, the first step includes:

[0026] (1) Establish the centroid kinematic model, that is, the kinematic model of the hypersonic vehicle, and express the relevant state variables of the centroid motion in the form of a differential equation system as the specific form of the differential terms of the centroid motion state variables in the subsequent formulas;

[0027] ;

[0028] Among them, , and are the three-axis position coordinates of the hypersonic vehicle in the inertial coordinate system, with initial values of 0m, 18000m, and 0m respectively. , and are respectively , and 's first derivatives; is the flight path inclination angle of the vehicle, with an initial value of -20°, is the flight path deviation angle of the vehicle, with an initial value of 0°, and are respectively and 's first derivatives; is the mass of the vehicle, which is 1800m / s, is the acceleration of the vehicle in the flight path coordinate system. , , are respectively the projections of the acceleration along the X-axis, Y-axis, and Z-axis in the flight path coordinate system.

[0029] The guidance-related models in the line-of-sight coordinate system are as follows:

[0030] ;

[0031] Among them, is the missile-to-target distance, with an initial value of 28913.7m, is 's first derivative, is 's second derivative; is the line-of-sight inclination angle, with an initial value of -38.5°, is the line-of-sight deviation angle, with an initial value of -45°, and are respectively and 's first derivatives, is 's second derivative; is the acceleration of the vehicle in the line-of-sight coordinate system. , , are respectively the projections of the acceleration along the X-axis, Y-axis, and Z-axis in the line-of-sight coordinate system.

[0032] and 's conversion relationship is as follows:

[0033] ;

[0034] ;

[0035] Among them, represents the uncertainty generated by the change in the aircraft speed on the Y-axis and Z-axis. The superscript T represents the transpose of the matrix. The matrix is an intermediate variable, and all its elements are intermediate variables.

[0036] The normal acceleration of the aircraft in the flight path coordinate system can be expressed by aerodynamic force and attitude angle:

[0037] ;

[0038] ;

[0039] Among them, represents the lift force, represents the side force, (or M) represents the mass of the aircraft changing with time, with an initial value of 1200 kg, and t represents the time, is the gravitational acceleration, with a value of 9.8 m / s 2 , is the angle of attack, with an initial value of 9.92°, is the sideslip angle, with an initial value of 0.81°, is the velocity inclination angle, with an initial value of 0.86°, , , and are aerodynamic parameters, which are 57.15, -0.081, 0.09, -56.32 respectively; , respectively represent the uncertainty of the aerodynamic parameters in the Y-axis direction and the uncertainty of the aerodynamic parameters in the Z-axis direction in the flight path coordinate system, , , t is the time, is the dynamic pressure, is the air density at the altitude where the aircraft is located, , is the characteristic area. Preferably, .

[0040] (2) Establish the dynamic model as:

[0041] ;

[0042] ;

[0043] ;

[0044] ;

[0045] ;

[0046] ;

[0047] ;

[0048] wherein, is the mass of the aircraft, is the acceleration due to gravity, , and are the first derivatives of , and respectively, , , are the components of the angular velocity of the body coordinate system relative to the ground coordinate system on the respective axes of the body coordinate system, and the initial values are 0.1 rad / s, 0.1 rad / s, and 0.1 rad / s respectively, , , are the first derivatives of , , respectively; , , , , , are unknown uncertainties caused by factors such as external disturbances, parameter perturbations, or unmodeled dynamics; , , are aerodynamic moments, , , are the rolling moment, yaw moment, and pitch moment generated by mass change; , , , , , , are aerodynamic parameters, which are 0.45, -0.38, 2.13, -27.3, -26.6, -28.15, -27.9 respectively; , , are control surfaces; is the reference length, preferably 0.69; , , are the moments of inertia about the three axes, which are 100, 5800, and 5700 respectively.

[0049] Specifically, in the second step, the integrated guidance and control model in strict feedback form is established based on the kinematic and dynamic models as follows:

[0050] ;

[0051] where, , , , are the model states, and , , , are respectively the first derivatives of , , , . represents the vector composed of line-of-sight angles, represents the vector composed of line-of-sight angular rates, represents the vector composed of attitude angles, represents the vector composed of angle of attack and sideslip angle, represents the vector composed of three-axis angular velocities, , , and are aerodynamic parameters, , , are the unknown disturbances and parameter uncertainties in the model;

[0052] ;

[0053] ;

[0054] ;

[0055] ;

[0056] ;

[0057] ;

[0058] ;

[0059] ;

[0060] ;

[0061] ;

[0062] ;

[0063] ;

[0064] Among them, represents the control input, which is a vector composed of three-axis control surfaces , , ; , , , , , are the non-linear terms in the strict feedback model; , are respectively the projections of the acceleration along the Y-axis and Z-axis in the line-of-sight coordinate system, , , are the unknown disturbances and parameter uncertainties in the model; and are respectively the lift force excluding the uncertainty term and the side force excluding the uncertainty term. diag represents a diagonal matrix, , are intermediate variables, l represents the reference length, , , represent the reciprocals of the three-axis moments of inertia, , , represent aerodynamic parameters, is an intermediate variable.

[0065] Specifically, the third step includes:

[0066] Design the first tracking error :

[0067] ;

[0068] Among them, is the desired line-of-sight angle, is the desired line-of-sight tilt angle, is the desired line-of-sight deviation angle.

[0069] Design the virtual control signal :

[0070] ;

[0071] Among them, , are design parameters.

[0072] Design the first filter:

[0073] ;

[0074] Among them, is the filter parameter, is the filter state, is the first derivative of.

[0075] Design the second tracking error :

[0076] ;

[0077] Let ;

[0078] Among them, is the infimum of, is the reciprocal of, is multiplied by the supremum of the modulus of the interference . Denotes the infimum, Denotes the supremum, Denotes the modulus of the vector.

[0079] Design the second adaptive law:

[0080] ;

[0081] ;

[0082] Among them, is the estimate of, is the first derivative of, is the estimate of, is the first derivative of, is an intermediate variable, , , , are design parameters.

[0083] Design the virtual control signal :

[0084] ;

[0085] Among them, is an intermediate variable, , , are design parameters.

[0086] Design the second filter:

[0087] ;

[0088] wherein, is the filter state, is the first derivative of, is the design parameter.

[0089] Design the third tracking error :

[0090] ;

[0091] wherein, is the tracking signal of.

[0092] Let ;

[0093] wherein, is the supremum of the modulus of the disturbance.

[0094] Design the third adaptation law:

[0095] ;

[0096] wherein, is the estimation of, is the first derivative of, , are the design parameters.

[0097] Design the virtual control signal :

[0098] ;

[0099] wherein, , are intermediate variables, , , , are the design parameters.

[0100] Design the third filter:

[0101] ;

[0102] wherein, is the filter parameter, is the filter state, is The first derivative.

[0103] Design the fourth tracking error :

[0104] ;

[0105] Let ,

[0106] ;

[0107] where \(l\) represents the reference length, is the infimum of the elements in is the reciprocal of is the product of and the supremum of the norm of the disturbance is the estimate of is the estimate of , , denotes the absolute value.

[0108] Let ;

[0109] where is the sign function, when , , when , .

[0110] Design the intermediate variable , where is the design parameter.

[0111] Design the fourth adaptation law as follows:

[0112] ;

[0113] ;

[0114] where , , , are the design parameters.

[0115] Design the control input as follows:

[0116] ;

[0117] where , are design parameters.

[0118] By using the method of the present invention for the integrated design of hypersonic vehicle guidance and control, the vehicle can hit a stationary target at a desired angle in the presence of time-varying inertial parameters, parameter uncertainties, and external disturbances, improving the anti-interference ability of the system.

[0119] Using the present invention to conduct a simulation experiment according to the above design, the desired line-of-sight angle is designed as , .

[0120] Figure 3 is the flight trajectory of the hypersonic vehicle, Figure 4 is the distance between the hypersonic vehicle and the target. It can be seen that the vehicle can hit the stationary target; Figure 5 is the line-of-sight angle of the hypersonic vehicle, Figure 5 where a is the tracking effect of the line-of-sight inclination angle, Figure 5 where b is the tracking effect of the line-of-sight deviation angle. It can be seen that the vehicle's line-of-sight angle can converge to the desired line-of-sight angle and attack the target at the desired line-of-sight angle; Figure 6 is the attitude angle of the hypersonic vehicle, Figure 6 where a is the angle-of-attack curve, Figure 6 where b is the sideslip-angle curve, Figure 6 where c is the velocity-inclination-angle curve. It can be seen that all attitude angles are bounded and converge to near 0.

[0121] The content not described in detail in the specification of the present invention belongs to the prior art well-known to those skilled in the art.

Claims

1. An adaptive anti-interference integrated guidance and control method for hypersonic vehicles, characterized in that It includes the following steps: Step 1: Establish the kinematic model and dynamic model of a hypersonic vehicle with time-varying inertial parameters; Step 2: Separate the time-varying inertial parameters, and based on the kinematic model and dynamic model of the hypersonic vehicle, establish an integrated guidance and control model in strict-feedback form; Step 3: Design an adaptive law to estimate the inertial parameters of the vehicle and the bounds of the unknown disturbances, and based on the dynamic surface method, combine the adaptive law to design an integrated guidance and control scheme to complete the integrated method for adaptive anti-disturbance guidance and control of the hypersonic vehicle.

2. An adaptive anti-interference guidance and control integrated method for a hypersonic vehicle according to claim 1, characterized in that The first step includes the following steps: Step 1: Establish a kinematic model of the hypersonic vehicle, i.e., the kinematic model of the center of mass, and represent the relevant state variables of the center of mass motion in the form of a differential equation system , where , and are the position coordinates of the hypersonic vehicle in the inertial coordinate system, , and are the first derivatives of , and respectively; is the flight path angle of the vehicle, is the flight path deviation angle of the vehicle, and are the first derivatives of and respectively, and the superscript T represents the transpose of the matrix; the input is the acceleration of the vehicle in the flight path coordinate system, , , are the projections of the acceleration of the vehicle in the flight path coordinate system along the X-axis, Y-axis, and Z-axis in the flight path coordinate system respectively; the guidance-related model in the line-of-sight coordinate system is , and the input is the acceleration of the vehicle in the line-of-sight coordinate system, , , are the projections of the acceleration of the vehicle in the line-of-sight coordinate system along the X-axis, Y-axis, and Z-axis in the line-of-sight coordinate system respectively; and have a conversion relationship; is the missile-to-target distance, is the second derivative of ; is the line-of-sight angle, is the line-of-sight deviation angle, and are the second derivatives of and the second derivative of respectively; the projection of the acceleration of the vehicle in the flight path coordinate system along the Y-axis and the projection of the acceleration of the vehicle in the flight path coordinate system along the Z-axis are determined by the lift , side force , angle of attack , sideslip angle , and velocity inclination angle and track angle indicates; Step 2: Establish the kinetic model as , is the control input and is a vector composed of control surfaces. Among them, , and are the first derivatives of the angle of attack , sideslip angle and velocity tilt angle respectively. , , are the components of the rotational angular velocity of the body coordinate system relative to the ground coordinate system on each axis of the body coordinate system. , , are the first derivatives of , , respectively. , , are control surfaces.​​​​​​​​ 3. An integrated guidance and control method for an adaptive anti-interference hypersonic vehicle according to claim 2, characterized in that In the second step, the established integrated guidance and control model in strict-feedback form is: ； Among them, , , , are the model states, , , , are respectively the first-order derivatives of the model states , , , . Among them, represents the vector composed of line-of-sight angles, represents the vector composed of line-of-sight angular rates, represents the vector composed of attitude angles, represents the vector composed of the angle of attack , the sideslip angle ; represents the vector composed of three-axis angular velocities; is the mass of the aircraft, and t represents time; , , , , , are the nonlinear terms in the strict feedback model; , , are the unknown disturbances and parameter uncertainties in the model.

4. An adaptive anti-interference guidance and control integration method for a hypersonic vehicle according to claim 3, characterized in that, The third step includes: Design the first tracking error , the first tracking error is the difference between the vector formed by the line-of-sight angle and . is the desired line-of-sight angle, is the desired line-of-sight inclination angle, is the desired line-of-sight declination angle; Design the virtual control signal , design a filter to obtain the filter state of the virtual control signal and its first derivative with respect to time .

5. An integrated method for adaptive anti-interference guidance and control of a hypersonic vehicle according to claim 4, characterized in that, The third step also includes: Design the second tracking error , the second tracking error is the vector formed by the line-of-sight angular rate and the filter state . Let be 's infimum, be 's reciprocal, be 's product with the supremum of the modulus of the unknown disturbances and parameter uncertainties in the model 6. An integrated method for adaptive anti-interference guidance and control of a hypersonic vehicle according to claim 5, characterized in that, The third step also includes: Design the second adaptive law as: ； ； Among them, is the estimate of, is the first derivative of, is the estimate of, is the first derivative of, is an intermediate variable, , , , are design parameters; represents the modulus of a vector; Design a virtual control signal , and design a first-order filter to obtain the filter state of the virtual control signal and its first derivative with respect to time .

7. An adaptive anti-interference guidance and control integration method for a hypersonic vehicle according to claim 6, characterized in that The third step also includes: Design the third tracking error , the third tracking error is the vector formed by the attitude angles and the difference of, is the designed tracking signal; Let be the unknown disturbance and parameter uncertainty term in the model , and design the third adaptive law as follows: ； Among them, is the estimate of, is the first derivative of, , are design parameters; Design a virtual control signal , and design a filter to obtain the virtual control signal filter state and its first derivative with respect to time .

8. An adaptive anti-interference guidance and control integrated method for a hypersonic vehicle according to claim 7, characterized in that The third step also includes: Design the fourth tracking error , the fourth tracking error is the vector composed of the three-axis angular velocities and the filter state . Let be the infimum of the elements in , and be the reciprocal of . Let be the product of and the supremum of the modulus of the unknown disturbances and parameter uncertainties in the model; be 's estimate, and be 's estimate. Let: ； Among them, is the sign function. When is the case, and when is the case, , where x represents a variable; l represents a reference length, , , represent the reciprocals of the three-axis moments of inertia; is the dynamic pressure, is the characteristic area, , , represent aerodynamic parameters.

9. An adaptive anti-interference integrated guidance and control method for a hypersonic vehicle according to claim 8, characterized in that, The third step also includes: Design the fourth adaptive law as follows: ； ； Among them, , , , are design parameters, denotes the first derivative of.

10. An adaptive anti-interference guidance and control integrated method for a hypersonic vehicle according to claim 9, characterized in that, The third step also includes: Design control input As follows: ； Among them, is a design parameter, , is a design parameter.

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