Aircraft guidance and control integrated design method based on control barrier function
By adopting an integrated guidance and control design method based on control obstacle functions, the problems of strong coupling and complex constraints in the glide phase of hypersonic vehicles were solved, achieving stable control of the vehicle and precise target arrival in the glide phase, thus improving the stability and safety of the system.
Patent Information
- Application Number
- CN202511606870.8
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-11-05
- Publication Date
- 2026-01-23
AI Technical Summary
The guidance and control design of hypersonic vehicles in the glide phase involves strong coupling, nonlinearity, and complex constraints. Existing technologies struggle to effectively address these factors, leading to design difficulties and system instability.
A guidance and control integrated design method based on the control obstacle function is adopted. By constructing the dynamics and kinematic equations of the hypersonic vehicle, and combining the obstacle Lyapunov function and dynamic surface control method, an integrated guidance and control model is designed to obtain the desired angle of attack, velocity tilt angle and track yaw angle commands, so as to achieve constraint and stable control of the vehicle's state variables.
This achievement enables hypersonic vehicles to maintain stability and accurately reach target positions under various process constraints, thereby improving the system's safety and stability.
Smart Images
Figure CN121386822A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The application belongs to the technical field of aircraft control, and particularly relates to an aircraft guidance control integrated design method based on a control barrier function. BACKGROUND
[0002] Due to the characteristics of high dynamic, complex multi-constraint, strong maneuverability and uncertainty of a hypersonic aircraft, compared with a traditional design method, an integrated design method considers the coupling relationship between a guidance system and a control system, can reduce design cost, and avoids a repeated design problem. A reentry glide phase is a key to enable application of hypersonic flight technology, but the flight environment of the phase is complex, the aircraft needs to reach a preset space region under the premise of meeting process constraints such as dynamic pressure, heat flux density and overload, and meet terminal constraints such as position, height and speed, and these factors all put forward strict requirements for a guidance and control method.
[0003] Compared with a traditional decoupling design model between channels, a guidance control integrated model has characteristics such as strong coupling and strong nonlinearity, which brings greater difficulty to control system design. On the other hand, in the actual design process of the guidance and control system, how to effectively deal with the influence of these constraints on the system is a problem that cannot be avoided. Therefore, the guidance and control integrated design considering process constraints is a more challenging problem, and there is less relevant technology disclosure so far.
[0004] Therefore, the application provides a hypersonic aircraft guidance control integrated design method based on a control barrier function. SUMMARY
[0005] The application aims to provide a hypersonic aircraft guidance control integrated design method based on a control barrier function to solve the problems in the background.
[0006] To achieve the above-mentioned purpose, the application provides the following technical scheme: a hypersonic aircraft guidance control integrated design method based on a control barrier function, comprising the following steps: S1: constructing hypersonic aircraft dynamics and kinematics equations based on Newton-Euler equations and coordinate system transformation; S2: constructing a hypersonic aircraft glide phase guidance control integrated model based on the center of mass dynamics, kinematics equations, center of mass dynamics, kinematics equations and an aerodynamic parameter model of the hypersonic aircraft; S3: obtaining an expected angle of attack command limit range by solving based on process constraints that need to be followed in the flight process; S4: taking a relative position of a target and the hypersonic aircraft as an input of the guidance control integrated model, obtaining expected speed dip angle and track yaw angle commands based on a guidance strategy of longitudinal range control and lateral azimuth error control; S5: Based on barrier Lyapunov function combined with dynamic surface control method, a hypersonic vehicle glide phase guidance control integrated control scheme is designed, and the expected dynamic of the mass center, the expected virtual command of the airflow system and the three-channel angular rate are established and tracked one by one. In order to solve the problem of reaching the specified terminal position while meeting the process constraints during the flight in the glide phase, the dynamic command limiting range is introduced to combine the barrier Lyapunov function to constrain the angle of attack command. S6: Combined with the around mass center dynamics equation of the aircraft, the control rudder deflection angle command of the aircraft is solved, which is input into the six-degree-of-freedom motion model of the hypersonic vehicle, so that the hypersonic vehicle accurately reaches the target position and completes the flight mission.
[0007] Preferably, in the S1 preparation process, the dynamics and kinematics equations of the hypersonic vehicle are as follows:
[0008] Among them, for the mass center dynamics equation in the ballistic coordinate system, is the velocity , the velocity inclination and the track yaw angle vector, for the mass center kinematics equation in the ballistic coordinate system, represents the geocentric distance , the latitude and the longitude vector , for the around mass center kinematics equation in the body coordinate system, represents the angle of attack , the sideslip angle and the tilt angle vector, for the around mass center dynamics equation in the body coordinate system, represents the roll angular velocity , the yaw angular velocity and the pitch angular velocity vector.
[0009] Preferably, the hypersonic vehicle guidance control integrated model constructed in the S2 preparation process is as follows:
[0010] Among them, each state vector is defined as: , , , , , The roll rudder deflection angle, the yaw rudder deflection angle and the pitch rudder deflection angle vector respectively; in the above formula, each nonlinear equation expression is as follows:
[0011] wherein, denotes the mass of the hypersonic vehicle, denotes the reference area of the vehicle, denotes the lift coefficient constant, denotes the dynamic pressure acting on the vehicle, which is expressed as , is the atmospheric density; denotes the gravity vector, denotes the inertial force vector caused by the earth rotation; is the characteristic length of the vehicle, are the roll moment of inertia, the yaw moment of inertia and the pitch moment of inertia, respectively, are the partial derivatives of the roll moment coefficient, the yaw moment coefficient and the pitch moment coefficient with respect to the sideslip angle and the partial derivative of the pitch moment coefficient with respect to the pitch deflection angle , respectively;
[0012] wherein, denotes the partial derivative of the lift coefficient with respect to the attack angle, and denote the partial derivatives of the roll moment coefficient with respect to and , respectively, and denote the partial derivatives of the yaw moment coefficient with respect to and , respectively, and denotes the partial derivative of the pitch moment coefficient with respect to the pitch deflection angle ;
[0013] wherein, is the uncertainty term of the lift coefficient, wherein, are the uncertainty terms of the roll moment coefficient, the yaw moment coefficient and the pitch moment coefficient, respectively.
[0014] Preferably, the expected angle of attack command range constructed by solving in the S3 is
[0015] wherein are the maximum values of the dynamic pressure, the heat flux density and the overload, respectively, and the maximum limit value of the dynamic pressure and the expected angle of attack command limit value can be obtained by solving according to the algebraic relationship between the vectors.
[0016] Preferably, the expected speed inclination angle command and the track yaw angle command obtained in the S4 are:
[0017] wherein, are the geocentric distance, latitude and longitude of the given desired target respectively.
[0018] Preferably, in the S5, the hypersonic vehicle glide phase guidance and control integrated control law is designed based on barrier Lyapunov function combined with dynamic surface control method as follows:
[0019] wherein, vector and are the sliding mode surface vectors designed based on barrier Lyapunov function combined with dynamic surface control, is an element in , is an element in , is an element in , and and are control parameter matrices, and are all positive diagonal matrices, are all normal numbers.
[0020] Preferably, the virtual control law constructed needs to pass through a filter as follows:
[0021] wherein, and are filter time constants; The final control law based on barrier Lyapunov function combined with dynamic surface control method in the S5 is shown as follows: .
[0022] Compared with the prior art, the present application has the beneficial effects that: The guidance and control integrated method proposed in the present application fully excavates the internal relationship between the guidance subsystem and the attitude control subsystem, and establishes a full state coupling guidance and control integrated design model; The present application considers the process constraints that must be obeyed by the hypersonic vehicle system in the glide phase, and through solving the algebraic relationship among the dynamic pressure, heat flow density and overload, the expected angle of attack instruction limit range is designed to constrain the vehicle state quantity, so as to ensure the safety performance of the vehicle and improve the stability of the system; The controller designed in the present application is designed based on barrier Lyapunov function combined with dynamic surface control method, and the control scheme can overcome the difficulties such as strong coupling, nonlinearity, uncertainty and strong constraints existing in the hypersonic vehicle system, and ensure that the system can accurately reach the target position under the condition of meeting multiple constraints. Attached Figure Description
[0023] Figure 1 This is a schematic diagram illustrating the principle of the integrated guidance and control scheme based on the control barrier function provided by this invention.
[0024] Figure 2 The curves show the changes in speed, velocity tilt angle, and yaw angle of the hypersonic vehicle in the simulation experiment.
[0025] Figure 3 The curves show the changes in angle of attack, sideslip angle, and tilt angle of a hypersonic vehicle during the simulation experiment.
[0026] Figure 4 The figures show the three-channel angular velocity variation curves of a hypersonic vehicle in a simulation experiment.
[0027] Figure 5 The curves showing the change in the deflection angle of the hypersonic vehicle during the simulation experiment are shown.
[0028] Figure 6 This is a three-dimensional trajectory diagram of a hypersonic vehicle in a simulation experiment.
[0029] Figure 7 This is a two-dimensional trajectory diagram of a hypersonic vehicle in a simulation experiment. Detailed Implementation
[0030] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.
[0031] Figure 1 This is a schematic diagram of the control scheme in this embodiment. As shown in the figure, the state of the hypersonic vehicle relative to the ground system is first measured by sensors; then, the obtained state variables of the vehicle, the acquired position information of the target relative to the ground system, and the set control parameters are substituted into the designed integrated guidance and control scheme to calculate the rudder deflection angle.
[0032] In this embodiment, based on the integrated guidance and control model of a hypersonic vehicle during the glide phase, an integrated guidance and control method for hypersonic vehicles based on a control obstacle function is designed, including the following steps: Step one: Based on Newton-Euler equation and coordinate transformation, the dynamics and kinematics equations of hypersonic vehicle are constructed. According to the model of hypersonic vehicle, the six-degree-of-freedom nonlinear model of vehicle is established, including: the equations of mass center dynamics and kinematics in trajectory coordinate system, the equations of mass center dynamics and kinematics in body coordinate system. The above six-degree-of-freedom motion equations of vehicle can be obtained from Chapter 5 of Hypersonic Vehicle Guidance and Control Integration Design.
[0033] Step two: The hypersonic vehicle glide phase guidance and control integration model is established. According to the six-degree-of-freedom nonlinear model of vehicle and the linear equation of aerodynamic parameters, the hypersonic vehicle glide phase guidance and control integration model is established, and the state variables are defined as follows: the velocity inclination angle and the track yaw angle vector , the angle of attack, the sideslip angle and the tilt angle vector , the angle of attack and the tilt angle related vector , the three-channel angular velocity vector , the rudder angle vector .
[0034] The constructed hypersonic vehicle guidance and control integration model is as follows:
[0035] In the formula, the velocity inclination angle and the track yaw angle are represented by and respectively, and the angle of attack, the sideslip angle and the tilt angle of the vehicle are represented by , and respectively, the roll, yaw and pitch angular velocities are respectively, and the left elevator, right elevator and rudder are represented by . All the above parameters are given by the outside.
[0036] The expressions of each nonlinear equation in formula (11) are as follows:
[0037] In the formula, represents the mass of the hypersonic vehicle, represents the reference area of the vehicle, represents the lift coefficient constant, represents the dynamic pressure acting on the vehicle, and its expression is , is the atmospheric density; represents the gravity vector, represents the inertial force vector caused by the earth rotation; is the characteristic length of the vehicle, are the roll, yaw and pitch moment of inertia respectively, are the partial derivatives of the roll, yaw and pitch moment coefficients with respect to the sideslip angle and the partial derivatives of the pitch moment coefficient with respect to the angle of attack respectively. All the above parameters are externally input.
[0038]
[0039] where, is the partial derivative of the lift coefficient with respect to the angle of attack, and are the partial derivatives of the roll moment coefficient with respect to and respectively, and are the partial derivatives of the yaw moment coefficient with respect to and respectively, and is the partial derivative of the pitch moment coefficient with respect to the elevator deflection angle . All the above parameters are externally input.
[0040]
[0041] where, is the uncertainty term of the lift coefficient. Wherein, are the uncertainty terms of the roll, yaw and pitch moment coefficients respectively. All the above parameters are externally input.
[0042] Step three: during the gliding flight of the hypersonic vehicle, a series of complex constraints are imposed. Based on the process constraints that need to be followed during the flight, the expected angle of attack command limit range is obtained by solving. The heat flux density , dynamic pressure and overload need to be satisfied as follows:
[0043] By solving the algebraic relationship between the dynamic pressure, heat flux density and overload, the expected angle of attack command limit range is designed to constrain the state quantity of the vehicle. The expected angle of attack command limit range is
[0044] where are the maximum values of the dynamic pressure, heat flux density and overload respectively. According to the algebraic relationship between the vectors, the maximum limit value of the dynamic pressure and the expected angle of attack command limit value can be obtained by solving.
[0045] Step four: Obtain the desired velocity pitch angle and track yaw angle command. Based on the guidance strategy of the range and azimuth error control, the control tasks of the longitudinal and lateral flight channels are realized, and then the desired velocity pitch angle and track yaw angle command are obtained .
[0046] Based on the quasi-equilibrium glide condition and the spherical geometry theory, the velocity pitch angle and track yaw angle command are:
[0047] wherein, is the current geodetic distance of the hypersonic vehicle, which can be obtained by external sensors, the latitude and longitude of the hypersonic vehicle at the current time are represented as and , the latitude and longitude information is obtained by external sensors; is the geodetic distance of the given target, and are the latitude and longitude of the given desired target position, respectively.
[0048] Step five: Design the hypersonic vehicle glide segment guidance control integration control scheme based on barrier Lyapunov function combined with dynamic surface control method. Through the guidance control integration model established in step two and the guidance strategy designed in step four, the guidance control integration control law of the hypersonic vehicle is designed, and the angle of attack command restriction range designed in step three is used to constrain the state quantity of the vehicle. The specific structure of the vehicle system is shown in Figure 1 .
[0049] Based on barrier Lyapunov function combined with dynamic surface control method, the guidance control integration control scheme of hypersonic vehicle glide segment is designed. The designed control scheme is: the dynamic surface control method is used to design the guidance law and attitude control law, and the barrier Lyapunov function and the angle of attack command restriction range are used in the attitude control law part to ensure that the constrained state of the closed-loop system meets the constraint condition requirement. The specific design process is as follows: By tracking the desired velocity pitch angle and track yaw angle, the guidance law is designed. The designed dynamic surface vector is described as follows:
[0050] wherein, represent the elements in the two-dimensional dynamic surface vector . Taking the first order derivative of formula (18) with respect to time and connecting the whole state coupling guidance control integration design model, the following formula can be obtained:
[0051] will , The control law is designed as virtual control input, and the control law is designed as follows according to the reachable condition:
[0052] wherein, is a positive parameter matrix to be adjusted, is a positive constant to be designed. and are the required attack angle and the required tilt angle respectively, the attack angle command and the tilt angle command can be determined according to formula (20). Assuming that the BTT turning mode is adopted by the hypersonic vehicle, the command of the sideslip angle satisfies the condition . Combined with the range of the angle of attack command designed in step three, the expected airflow angle can be obtained as follows:
[0053] wherein the saturation function is , a virtual control law is constructed, and it is passed through the filter as follows:
[0054] wherein, is the filter time constant to be designed.
[0055] In the design of the attitude control law, the tracking of the attack angle command, the sideslip angle command and the tilt angle command is performed. Then, the required rudder deflection angle of the control surface will be calculated according to these commands. The purpose of the attitude control is to generate the necessary rudder deflection angle command, which includes the roll rudder deflection angle, the yaw rudder deflection angle and the pitch rudder deflection angle. The obtained control surface rudder deflection must ensure that the actual airflow angle is consistent with the required airflow angle generated by the guidance law. is the reference tracking variable of the attitude control, so the designed dynamic surface vector is described as follows:
[0056] wherein, is an element in . The first-order derivative of formula (23) with respect to time is obtained, and the integrated design model of the full state coupled guidance control is obtained as follows:
[0057] The virtual control law designed according to the reachable condition is as follows:
[0058] wherein, is a diagonal matrix of feedback coefficients to be designed, is a positive constant to be designed, is the vector of the barrier Lyapunov function designed approaching law, .
[0059] Construct a virtual control law and pass it through the filter as follows
[0060] where, is the designed filter time constant.
[0061] The desired command as the virtual control input, so the designed dynamic surface vector is described as follows:
[0062] where, is the element in Take the first-order derivative of formula (23) with respect to time and integrate the full state coupling guidance control integrated design model, and obtain
[0063] The virtual control law designed according to the attainable condition is:
[0064] where, is the feedback coefficient diagonal matrix to be designed, is the designed normal number.
[0065] Therefore, according to formula (20), formula (25 and (29), the control law based on the barrier Lyapunov function combined with the dynamic surface control method proposed is as follows:
[0066] The following is based on the above-mentioned hypersonic vehicle glide segment guidance and control integrated design method for simulation verification: Table 1 records the initial conditions of the hypersonic vehicle, including initial position, initial state quantity and initial rudder angle, etc. The terminal target is designed to be static, and the position of the target is: height 40 km, longitude-24°, latitude 50°; and each rudder angle is limited within 30 degrees, that is, The uncertainty of the vehicle mass center dynamics model is set as follows: The parameters required for the dynamic surface controller based on the barrier Lyapunov function are shown in formula (31): Table 1. Initial conditions of hypersonic vehicle
[0067]
[0068] Based on the initial conditions provided in Table 1, the proposed control scheme was simulated and tested. Specific simulation results are as follows: Figures 2-7 As shown. The aircraft's speed, velocity tilt angle, and track yaw angle are as follows: Figure 2 As shown in the figure, it can be clearly seen that the range of the flight heading angle is very small, varying between 0 degrees and -1 degrees most of the time. Meanwhile, the curves for the angle of attack, sideslip angle, and roll angle of a hypersonic vehicle are as follows: Figure 3 As shown. From Figure 3 It can be observed that all state variables are smooth and bounded, while simultaneously satisfying the limits of the angle of attack command. And the angular velocity... The curve is as follows Figure 4 As shown, the conclusion is similar to that of the airflow angle; all three angular velocities are smooth and bounded. The actual deflection angles of the left and right elevators and the rudder are as follows: Figure 5 As shown, according to Figure 5 As shown, All are bounded and smooth, satisfying the constraints. The three-dimensional and two-dimensional trajectories of the hypersonic vehicle are shown in the figures. Figure 6 and Figure 7 As shown.
[0069] Therefore, the state variables of the hypersonic vehicle are stable and controllable during the glide phase, effectively solving the problem of simultaneously satisfying process constraints when reaching the designated terminal position during the glide phase, thus ensuring system stability. Based on the above analysis, the effectiveness of the designed IGC scheme is proven. Although embodiments of the present invention have been shown and described in detail above, it will be understood by those skilled in the art that various changes, modifications, substitutions, and variations can be made to these embodiments without departing from the principles and spirit of the present invention, the scope of which is defined by the appended claims and their equivalents.
Claims
1. A method for integrated design of guidance and control of aircraft based on control barrier function, comprising, characterized in that: S1: Based on Newton Euler equation and coordinate transformation, the dynamics and kinematics equations of hypersonic vehicle are constructed; S2: Based on the center of mass dynamics and kinematics equations of hypersonic vehicle, around the center of mass dynamics and kinematics equations and the aerodynamic parameter model, the hypersonic vehicle glide segment guidance control integrated model is constructed; S3: Based on the process constraints that need to be followed during flight, the expected angle of attack command limit range is obtained by solving; S4: The relative position of the target and the hypersonic vehicle is taken as the input of the guidance control integrated model, and based on the guidance strategy of longitudinal range control and lateral azimuth error control, the expected speed dip angle and track yaw angle command are obtained; S5: Based on barrier Lyapunov function combined with dynamic surface control method, the hypersonic vehicle glide segment guidance control integrated control scheme is designed, and the expected virtual command of center of mass expected dynamics, airflow system and three channel angular rate is established and tracked in turn. In order to solve the problem of reaching the specified terminal position while meeting the process constraints during flight in glide segment, dynamic command limit range is introduced to combine barrier Lyapunov function to constrain the angle of attack command; S6: Combined with the center of mass dynamics equation of the vehicle, the vehicle control rudder deflection angle command is calculated, which is input into the six degree of freedom motion model of hypersonic vehicle, so that the hypersonic vehicle can accurately reach the target position and complete the flight task.
2. The control barrier function based integrated guidance and control design method for aircraft according to claim 1, wherein: In the preparation process of S1, the dynamics and kinematics equations of hypersonic vehicle are as follows: where, for the equations of motion of the center of mass in the ballistic coordinate system, is the velocity , the velocity inclination , and the track yaw angle vector, for the equations of motion of the center of mass in the ballistic coordinate system, denotes the geocentric distance , the latitude , and the longitude vector , for the equations of motion of the body about the center of mass in the body coordinate system, denotes the angle of attack , the angle of sideslip , and the angle of bank vector, for the equations of motion of the body about the center of mass in the body coordinate system, denotes the roll angular velocity , the yaw angular velocity , and the pitch angular velocity vector.
3. The control barrier function based integrated guidance and control design method for aircraft according to claim 1, wherein: In the preparation process of S2, the hypersonic vehicle guidance control integrated model constructed is as follows: where each state vector is defined as: , , , , , are the roll, yaw and pitch rudder deflection vectors, respectively; in the above equations, each non-linear equation is expressed as follows: wherein denotes the hypersonic vehicle mass, denotes the reference area of the vehicle, denotes the lift coefficient constant, denotes the dynamic pressure acting on the vehicle, which is expressed as , is the atmospheric density; denotes the gravity vector, denotes the inertial force vector caused by the earth rotation; is the characteristic length of the vehicle, are the roll, yaw and pitch moments of inertia, respectively, are the partial derivatives of the roll, yaw and pitch moment coefficients with respect to the sideslip angle and the partial derivatives of the pitch moment coefficient with respect to the angle of attack , respectively. in, This represents the partial derivative of the lift coefficient with respect to the angle of attack. and These represent the rolling moment coefficients with respect to... and The partial derivatives, and These represent the yaw moment coefficients with respect to... and The partial derivatives of, and This indicates that the pitch moment coefficient is related to the pitch deflection angle. The partial derivatives; wherein, is an uncertainty term for the lift coefficient. Wherein, are uncertainty terms for the roll, yaw and pitch moment coefficients, respectively.
4. The control barrier function based integrated guidance and control design method for aircraft according to claim 1, wherein: The expected angle of attack command limit range constructed by solving in S3 is wherein are the maximum values of dynamic pressure, heat flux density and overload, respectively, and the maximum limit value of dynamic pressure can be obtained by solving the algebraic relationship between the vectors and the desired angle of attack command limit value .
5. The control barrier function based integrated guidance and control design method for aircraft according to claim 1, wherein: The expected speed dip angle and track yaw angle command obtained in S4 are: wherein, respectively the geocentric distance, latitude and longitude of the given desired target.
6. The control barrier function based integrated guidance and control design method for aircraft according to claim 1, wherein: In S5, the control law of hypersonic vehicle glide segment guidance control integration designed based on barrier Lyapunov function combined with dynamic surface control method is: where the vector and are the sliding mode surface vectors designed based on the barrier Lyapunov function combined with the dynamic surface control, is the element in is the element in is the element in is the element in is the element in is the element in and are control parameter matrices, and are positive diagonal matrices, are positive constants.
7. The control barrier function based integrated guidance and control design method for aircraft according to claim 6, wherein: The filter that the virtual control law needs to pass through is as follows: wherein and are filter time constants; The control law finally obtained based on barrier Lyapunov function combined with dynamic surface control method in S5 is as follows: 。