Spacecraft attitude and orbit integrated finite time sliding mode control method and system

Abstract

The application provides a kind of aircraft attitude and orbit integrated finite time sliding mode control method and system, including establishing the relative attitude and orbit integrated kinematics and dynamics model of aircraft and target based on Lie group SE (3);Non-singular terminal sliding mode surface is set, and hyperbolic function is used instead of sign function to construct sliding mode surface expression;Combined with disturbance observer, set finite time sliding mode controller, wherein the disturbance observer is used to estimate system disturbance and generate disturbance estimate value;According to the sliding mode controller, control input is generated, and attitude and trajectory error finite time convergence is realized.The scheme of the application has fast convergence, high-precision tracking and strong anti-interference ability, and is suitable for precise control of high-speed maneuverable aircraft in complex space tasks.

Description

Technical Field

[0001] This invention relates to the field of aircraft control technology, specifically to an integrated modeling and control technology scheme for aircraft attitude and trajectory based on Lie group SE(3). Background Technology

[0002] In traditional aircraft control methods, attitude and trajectory are typically designed separately, leading to the neglect of dynamic coupling and resulting in problems such as response lag, insufficient control precision, and resource waste. This makes it difficult to meet the precise control requirements of high-speed maneuvering aircraft in modern aerospace missions, where attitude and trajectory are highly coupled. While vector algebra-based methods can describe attitude and trajectory separately, they are still essentially separate models and fail to truly solve the dynamic coupling problem. Dual quaternion-based methods can uniformly represent attitude, but parameter non-uniqueness and decoupling phenomena cause the model to fail under specific attitudes. Conformal geometric algebra-based methods, due to their high computational complexity, are difficult to meet real-time control requirements. Regarding control strategies, traditional back-thrust tracking control and asymptotically convergent sliding mode control can achieve stable tracking, but their convergence time depends on the initial state, making it difficult to guarantee the achievement of desired performance within a finite time. Furthermore, aircraft often face challenges such as aerodynamic disturbances and model uncertainties in actual missions. Existing control methods lack robustness in strong disturbance environments, easily leading to tracking error divergence or severe control input jitter, seriously affecting system reliability. Therefore, there is an urgent need in this field for a control method that can describe the attitude and trajectory coupling relationship globally without singularities, and has both fast convergence and strong anti-interference capabilities, in order to meet the precise control requirements of high-speed maneuvering aircraft in complex aerospace missions. Summary of the Invention

[0003] This invention provides a high-speed maneuvering aircraft attitude and orbit integrated modeling and finite-time sliding mode control technology based on Lie group SE(3), which solves the problems of insufficient dynamic coupling description, weak anti-interference ability and slow convergence speed in traditional methods.

[0004] This invention provides a finite-time sliding mode control method for integrated attitude and orbit control of an aircraft, comprising the following processes: A relative attitude and orbit integrated kinematic and dynamic model of the aircraft and the target is established based on the Lie group SE(3); Set a non-singular terminal sliding surface and use hyperbolic functions to replace the symbolic functions to construct the sliding surface expression; A finite-time sliding mode controller is configured in conjunction with an interference observer, wherein the interference observer is used to estimate system interference and generate interference estimates. The control input is generated based on the sliding mode controller to achieve convergence of attitude and trajectory errors within a finite time.

[0005] Furthermore, the setting of the non-singular terminal sliding surface includes constructing a sliding surface based on generalized velocity error and using a hyperbolic secant function to suppress chattering during the sliding mode control process. Moreover, the interference observer generates an estimate of the system interference and its derivative through multi-level correction terms, adjusts the observer output using nonlinear gain coefficients and adaptive parameter matrices, and then constructs the corresponding observer error equation.

[0006] Furthermore, the expression of the finite-time sliding mode controller includes a sliding surface term, a disturbance estimation compensation term for the disturbance observer, control coefficients, the minimum eigenvalue function of the system inertia tensor matrix, and a hyperbolic function term for smoothing the control input.

[0007] Moreover, the finite-time convergence is proven by Lyapunov stability theory, and the convergence time is determined by the initial state, sliding surface parameters, and control coefficients.

[0008] On the other hand, the present invention also provides an integrated attitude and orbit control system for aircraft, comprising the following modules: The kinematics and dynamics modeling module is used to construct an integrated attitude and orbit model of the aircraft and the target based on the Lie group SE(3); The interference observer module is used to estimate and output interference estimates; A non-singular terminal sliding mode controller module is used to construct a sliding mode surface using a hyperbolic function and generate control inputs based on the disturbance estimate. The stability analysis module is used to verify the finite-time convergence of the system state.

[0009] Furthermore, the non-singular terminal sliding mode controller module is configured to embed the hyperbolic secant function into the sliding surface derivative function to suppress chattering, adjust the estimation accuracy of the interference observer through a multi-level correction mechanism, and ensure finite-time convergence by combining control coefficients and eigenvalue functions.

[0010] On the other hand, the present invention also provides an electronic device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft as described above.

[0011] On the other hand, the present invention also provides a non-transitory computer-readable storage medium having a computer program stored thereon, wherein the computer program, when executed by a processor, implements the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft as described above.

[0012] On the other hand, the present invention also provides a computer program product, including a computer program that, when executed by a processor, implements the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft as described above.

[0013] The beneficial effects of this invention are: (1) This invention proposes a finite-time sliding mode control scheme for aircraft attitude and trajectory integration based on the Lie group SE(3). By modeling the relative attitude of the aircraft as elements in the Lie group SE(3), the coupling relationship between the aircraft attitude and trajectory can be accurately described, avoiding the problem of separate modeling in traditional methods. (2) By designing a non-singular terminal sliding surface and combining it with a disturbance observer, the robustness of the system in complex environments is effectively improved. Especially under strong disturbances and model uncertainties, the system can suppress the influence of external disturbances on the system in real time, avoiding the chattering phenomenon that may occur in traditional sliding mode control. (3) In addition, it can achieve accurate tracking of attitude and trajectory within a finite time, ensuring that the aircraft can quickly converge to the desired state, meeting the strict requirements of high-speed maneuvering aircraft for control accuracy. Attached Figure Description

[0014] Figure 1 This is a flowchart illustrating the attitude and trajectory integrated finite-time control method provided in an embodiment of the present invention; Figure 2 This is a graph showing the angular velocity error of the aircraft according to an embodiment of the present invention; Figure 3 This is the speed error curve of the aircraft in an embodiment of the present invention; Figure 4 This refers to the attitude tracking error of the aircraft in an embodiment of the present invention. Figure 5 This refers to the aircraft position tracking error in an embodiment of the present invention. Figure 6 This is the overload curve of the aircraft in an embodiment of the present invention; Figure 7 This is the relative motion curve between the aircraft and the target in an embodiment of the present invention. Detailed Implementation

[0015] To make the objectives, technical solutions, and advantages of this invention clearer, the technical solutions of this invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this invention. All other embodiments obtained by those skilled in the art based on the embodiments of this invention without creative effort are within the scope of protection of this invention.

[0016] To address the problems existing in the prior art, this invention establishes an integrated attitude and orbit model of a high-speed maneuvering aircraft based on the Lie group SE(3), and designs a finite-time sliding mode control scheme based on an interference observer.

[0017] To address the problems of insufficient control accuracy, response lag, and weak anti-interference capability caused by traditional separate modeling of aircraft attitude and trajectory, this invention models the relative attitude of the aircraft as elements in a Lie group SE(3). Considering model uncertainty, an integrated kinematic and dynamic model of the model and the target attitude and trajectory is established. A non-singular terminal sliding surface is designed, and a finite-time sliding mode controller is constructed in conjunction with an interference observer. The global convergence of attitude and trajectory errors of the closed-loop system is proven in a finite time using Lyapunov stability theory. The control system provided by this invention has fast convergence, high-precision tracking, and strong anti-interference capability, and is suitable for the precise control of high-speed maneuvering aircraft in complex aerospace missions.

[0018] See Figure 1 The embodiment provides a finite-time sliding mode control method for integrated attitude and orbit control of an aircraft, which includes the following steps: Step S1: Define the relative attitude of the aircraft as an element in the Lie group SE(3), establish a generalized relative velocity model, and derive the kinematic and dynamic equations of the relative attitude and orbit integration of the high-speed maneuvering aircraft based on the Lie group SE(3).

[0019] Lie groups SE(3) are mathematical structures that describe the motion of rigid bodies in three-dimensional space. Unlike traditional methods based on dual quaternions, this invention proposes to establish the attitude and orbit integrated kinematics and dynamic equations of an aircraft in Lie groups SE(3). The attitude of the aircraft is represented by rotation matrices and position vectors, providing an essential, concise and non-singular mathematical framework for the strong coupling control of attitude and orbit of high-speed maneuvering aircraft.

[0020] The preferred implementation method suggested in the embodiments includes the following sub-steps: Step S11: Establish an integrated kinematic and dynamic model of attitude and orbit for a single high-speed maneuvering aircraft.

[0021] The pose of the aircraft in the Lie group SE(3) It can be described as (1) in, It is a description of the aircraft's attitude. It is The orthogonal matrix is ​​the rotation matrix of the aircraft's body coordinate system relative to the inertial coordinate system, used to describe the aircraft's rotation in three-dimensional space. It is The column vectors represent the position of the aircraft's body coordinate system relative to the inertial coordinate system, and are used to describe the aircraft's translation in three-dimensional space. It is a three-dimensional special Euclidean group. It is a 1×3 zero matrix.

[0022] The generalized attitude velocity of an aircraft can be described as (2) in, A generalized description of the speed of an aircraft. Let be the angular velocity of the aircraft in its body coordinate system. Let be the linear velocity of the aircraft in its body coordinate system, with the superscript T indicating transpose. The second form of the generalized attitude-orbit velocity. for (3) Where the superscript × denotes the antisymmetric matrix operator, (3) is a special Euclidean group SE (3) The corresponding Lie algebra.

[0023] To establish an integrated attitude and orbit model of an aircraft, several adjoint operators need to be defined to describe the changes in the aircraft's position and velocity.

[0024] The adjoint transformation matrix of a rigid body in space It can be represented as (4) in, It is a 1×3 zero matrix.

[0025] The adjoint operator matrix of rigid body velocity It can be represented as (5) Its inverse adjoint operator It can be obtained through the dual operation of Lie algebras. (6) in, for The transpose of .

[0026] Based on the mapping relationship above, the integrated attitude and orbit kinematics and dynamics model of the aircraft in the inertial coordinate system is obtained. (7) in It is the inertial tensor matrix composed of the aircraft's mass and moment of inertia. For the gravity and gravitational torque of the aircraft, The aerodynamic forces and aerodynamic torques acting on the aircraft. Interference to the system For system control input, It is a position The time derivative of is the physical meaning of the instantaneous motion state of the aircraft on the SE(3) group.

[0027] For a certain axisymmetric X-type high-speed maneuvering aircraft, its aerodynamic coefficient model is as follows: (8) in, , , These represent the aerodynamic coefficients of the aircraft along the X, Y, and Z axes in the body coordinate system, respectively. It is the zero-lift drag coefficient, which is unrelated to lift. It is the induced drag coefficient, which depends on the magnitude of the lift. For the angle of attack of the aircraft, The sideslip angle of the aircraft. It is the angle of attack coefficient. It is the sideslip angle coefficient; , , It is the deflection angle of the control surface. It is the rudder efficiency coefficient. It is the elevator efficiency coefficient. , .

[0028] Its aerodynamic moment coefficient model is (9) In the formula, It is the rolling moment coefficient. It is the pitch moment coefficient. It is the yaw moment coefficient. The asymmetrical shape is caused by manufacturing errors. , , It is the deflection angle of the control surface. It is the rudder deflection rate; It is the dimensionless roll angular velocity in the x-direction. It is the dimensionless roll angular velocity in the y-direction. It is a dimensionless rate of change of sideslip angle. It is the dimensionless elevator deflection rate. , , , . , , These are the angular velocities of the aircraft in the three directions of the body coordinate system. It is the characteristic length of the aircraft. This represents the speed of the aircraft. For the angle of attack of the aircraft, The sideslip angle of the aircraft. It is the angle of attack rate. It is the roll derivative of the sideslip angle. It is the pitch derivative of the sideslip angle. It is the derivative of angle of attack and yaw. It is the aileron roll control derivative. It is the elevator roll cross derivative. It is the elevator pitch control derivative. It is the rudder yaw control derivative. It is the derivative of roll damping. It is the coupled effect of pitch angular velocity on roll moment. It is the pitch damping derivative. It is the derivative of the sideslip angular rate. It is the elevator pitch control derivative. It is the derivative of yaw damping. It is the dynamic effect of the rate of change of angle of attack on the yaw moment. It is the dynamic effect of rudder deflection rate on yaw moment.

[0029] The expressions for aerodynamic force and aerodynamic torque are as follows: (10) (11) In the formula It is the characteristic area. It is the feature length. It's air dynamic pressure. ,in It is the air density at the altitude of the aircraft. , , These are roll force, pitch force, and yaw force. , , These are rolling moment, pitching moment, and yaw moment.

[0030] Step S12: Establish an integrated kinematic and dynamic model of the relative attitude and orbit of the high-speed maneuvering aircraft and the target.

[0031] Let the desired pose of the target be... The aircraft's own attitude is Attitude and trajectory tracking error between the aircraft and the target for (12) (13) in, Let be the rotation error matrix of the two spacecraft. Here is the position error matrix for the two spacecraft. The transpose of the target spacecraft's attitude matrix. The tracking error is uniformly described using exponential coordinates on the Lie group SE(3) to represent the position of the target spacecraft.

[0032] (14) in This represents the tracking error in the attitude section. Indicates position tracking error. and It is expressed as follows (15) (16) (17) in, For rotation angle, The trace of the matrix, It is a 3×3 identity matrix. This is an extended form of the Rodriguez rotation formula.

[0033] Speed ​​tracking error It can be represented as (18) in, This is the transpose of the angular velocity error. For the transpose of the speed error, For Li Qun SE (3) is an accompanying representation. For the desired generalized speed.

[0034] Using exponential coordinates based on the Lie group SE(3) To describe the attitude tracking error of an aircraft, the kinematic equation for the attitude and trajectory error between the aircraft and the target can be expressed as: (19) In the formula (20) (twenty one) (twenty two) in, Exponential coordinates based on the Lie group SE(3) Time derivative, This is the velocity transformation matrix. For generalized velocity error, Used to handle rotational motion, Used to handle the coupling effect of rotation and translation. It is an identity matrix.

[0035] The relative attitude-orbit integrated dynamic equation based on the Lie group SE(3) can be described as follows: (twenty three) Step S2: Considering model uncertainties, establish an integrated kinematic and dynamic model of the relative attitude and orbit of the high-speed maneuvering aircraft and the target.

[0036] The present invention further proposes to design a non-singular terminal sliding surface and to use a hyperbolic function instead of a sign function to avoid oscillations on both sides of the sliding surface; Design an interference observer to estimate system interference and design control inputs based on the estimates.

[0037] For situations where model uncertainty exists, uncertainty modeling is introduced into the system. This indicates the uncertainty of the system model.

[0038] (twenty four) The system can be simplified to the following system model, that is, considering model uncertainties, the integrated attitude and orbit model of the aircraft and the target is: (25) in Indicates the defined part of the system. This is interference with the system.

[0039] (26) (27) (28) Step S3: Considering the nonlinear characteristics of the aircraft dynamics model and external disturbances, a nonsingular terminal sliding mode surface is constructed, and a finite-time sliding mode control algorithm is designed in conjunction with a disturbance observer. Specifically, a finite-time sliding mode controller is designed, where the disturbance observer is used to estimate system disturbances and generate disturbance estimates; control inputs are generated based on the sliding mode controller to achieve convergence of attitude and trajectory errors within a finite time.

[0040] This step, based on Lyapunov stability theory, proves that the system's state and error converge to zero in finite time.

[0041] In this embodiment, a fast non-singular terminal sliding surface is selected. The system state is determined after the sliding surface reaches zero. and All can converge to zero within a finite time. Due to the switching nature of sliding mode, chattering can occur, causing the system state to oscillate back and forth between the two sides of the sliding surface. In practical sliding surface design, non-singular terminal sliding mode controllers use hyperbolic functions instead of sign functions to reduce oscillations during sliding mode control.

[0042] Sliding surface As shown in formula (29).

[0043] (29) in, , These are the weighting coefficients. These are non-linear exponential coefficients.

[0044] Differentiating with respect to the sliding surface yields

[0045] (30) Here, sech() represents the hyperbolic secant function.

[0046] Substitute the system model (25) into equation (30). (31) Assume the disturbances to the system are bounded, and ,in, Denotes the i-th derivative. For system interference, An upper bound on the absolute value of the disturbance is provided for the corresponding threshold, which can be preset based on experiments or experience. An disturbance observer is designed to adjust the control input by estimating the error of the system disturbance, ensuring the robustness of the system. The disturbance observer generates an estimate of the system disturbance and its derivative through multi-level correction terms, adjusts the observer output using a nonlinear gain coefficient and an adaptive parameter matrix, and then constructs the observer error equation accordingly. (32) in, They are respectively The estimated value, , , for The corresponding time derivative, This is a first-level correction item. This is a second-level correction item. , , For nonlinear gain coefficients, For adaptive parameter matrix, It is a symbolic function.

[0047] The error equation of the observer can be obtained. (33) in, , , The derivative of the observation error, This is the derivative of the interference estimation error.

[0048] Based on the designed interference observer, the controller is designed as follows: (34) In the formula, . , For control coefficients, The expression for this finite-time sliding mode controller includes a sliding surface term, a disturbance estimation compensation term for the disturbance observer, control coefficients, the minimum eigenvalue function of the system inertia tensor matrix, and a hyperbolic function term for smoothing the control input.

[0049] When the sliding surface reaches After that, the system status and It can converge to zero in a finite amount of time; the convergence time is... for (35) in, This is the initial time.

[0050] Based on the above three steps, the entire attitude and orbit control process of the high-speed maneuvering aircraft is completed.

[0051] Based on the above embodiments, in order to verify the effectiveness of the control model and control algorithm for high-speed maneuvering aircraft established in this invention, the control system of high-speed maneuvering aircraft was integrated and designed in Matlab, and the following simulation verification was performed.

[0052] (1) Parameter settings mass of the aircraft Moment of inertia .

[0053] The initial attitude of the aircraft is .

[0054] The generalized initial velocity of the aircraft is .

[0055] The first three columns represent the aircraft's speed, and the last three columns represent the aircraft's angular velocity.

[0056] The target's pose is .

[0057] The generalized velocity of the target is .

[0058] For system model uncertainties, considering the perturbation range of the mass and inertia matrices, the following is... The system interference is Random numbers.

[0059] The parameters of the designed controller are shown in Table 1.

[0060] Table 1 Controller Parameters

[0061] in, It is a 6-dimensional identity matrix.

[0062] (2) Results Analysis Simulation results are as follows Figures 2-7 As shown. Figure 2 The angular velocity error curve of the aircraft. Figure 3 The aircraft speed error curve, Figure 4 For aircraft attitude tracking error, Figure 5 For aircraft position tracking error, Figure 6 The relative motion curves of the aircraft and the target. Figure 7 This is the overload curve for the aircraft.

[0063] Figure 2 The graph shows the angular velocity changes of the aircraft in three directions. It can be seen that the angular velocity changes rapidly initially, then gradually stabilizes at the target angular velocity. This indicates that the aircraft made significant attitude adjustments in the initial stage before entering a stable flight phase. Figure 4 It can be seen that the attitude tracking error of the aircraft is relatively large in the initial stage, but it will quickly converge to near zero. Figure 3 The graphs show the velocity changes of the aircraft in three directions. The aircraft initially has a high velocity, which then stabilizes within 20 seconds, maintaining the target velocity. From Figure 5 As can be seen, the aircraft's position error changes rapidly in the initial stage and converges to zero within approximately 40 seconds, indicating that the designed controller can effectively compensate for the initial position error. Figure 6 The aircraft's overload curve shows that it experienced significant overload in the initial stage, reaching a peak value of nearly 82, before rapidly decreasing and stabilizing around 40 seconds. The overload level remained low throughout the subsequent flight. This indicates that the aircraft underwent intense maneuvers during startup or attitude adjustment to quickly correct its flight direction and reach the predetermined speed, while the overload gradually decreased as flight stabilized.

[0064] Figure 7 illustrates the spatial relationship between the aircraft's trajectory (red line) and the target's trajectory (blue line) in three-dimensional space. As can be observed from the figure, the aircraft starts from its initial position (red dot), flies along a smooth curve under the control of the control system, and ultimately reaches the target position accurately. This demonstrates the effective target tracking capability of the designed controller in three-dimensional space.

[0065] The results above demonstrate that the proposed attitude and orbit integrated control method based on the Lie group SE(3) exhibits excellent performance in controlling the angular velocity, linear velocity, position error, and attitude error of high-speed maneuvering aircraft. The system can quickly suppress initial disturbances, achieve rapid error convergence, and maintain high accuracy in the steady-state phase. The robustness and dynamic performance of the control system have been fully verified, proving the applicability and effectiveness of the proposed method in complex environments.

[0066] In specific implementation, the method proposed in the technical solution of this invention can be automatically executed by those skilled in the art using computer software technology. System devices for implementing the method, such as computer-readable storage media storing the corresponding computer program of the technical solution of this invention and computer equipment including the computer program running the corresponding computer program, should also be within the protection scope of this invention.

[0067] The following embodiments describe the electronic device provided by the present invention. The electronic device described below can be referred to in correspondence with the finite-time sliding mode control method for integrated attitude and orbit control of aircraft described above.

[0068] The electronic device may include a processor, a communications interface, a memory, and a communication bus. The processor, communications interface, and memory communicate with each other via the communication bus. The processor can call logical instructions from the memory to execute the integrated attitude and orbit finite-time sliding mode control method for the aircraft, which mainly includes the software processing portion described above.

[0069] Furthermore, the logical instructions in the aforementioned memory can be implemented as software functional units and sold or used as independent products, and can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention, or the part that contributes to the prior art, or a part of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute all or part of the steps of the methods described in the various embodiments of the present invention. The aforementioned storage medium includes various media capable of storing program code, such as USB flash drives, portable hard drives, read-only memory (ROM), random access memory (RAM), magnetic disks, or optical disks.

[0070] In one possible embodiment, a high-speed maneuvering aircraft attitude and orbit integrated control system based on the Lie group SE(3) is provided, comprising: An integrated kinematic and dynamic model of an aircraft attitude and orbit is used to describe the relative attitude and orbit motion of the aircraft in the Lie group SE(3); An interference observer is used to estimate external interference and adjust the control strategy based on the estimate; A non-singular terminal sliding mode controller is used to design control inputs based on the aircraft's motion state, ensuring finite-time convergence of attitude and trajectory; A stability analysis module for analyzing the finite-time convergence of a system based on Lyapunov theory.

[0071] The designed controller is implemented as shown in formula (34) above. The non-singular terminal sliding mode controller module is configured to embed the hyperbolic secant function into the sliding surface derivative function to suppress chattering, adjust the estimation accuracy of the interference observer through a multi-level correction mechanism, and ensure finite-time convergence by combining control coefficients and eigenvalue functions.

[0072] On the other hand, embodiments of the present invention also provide a computer program product, the computer program product including a computer program, the computer program being stored on a non-transitory computer-readable storage medium, and when the computer program is executed by a processor, the computer is able to execute the software processing part of the integrated finite-time sliding mode control method for aircraft attitude and orbit provided by the above methods.

[0073] In another aspect, embodiments of the present invention also provide a non-transitory computer-readable storage medium storing a computer program thereon, which, when executed by a processor, implements the software processing portion of the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft provided by the above methods.

[0074] The device embodiments described above are merely illustrative. The units described as separate components may or may not be physically separate. The components shown as units may or may not be physical units; that is, they may be located in one place or distributed across multiple network units. Some or all of the modules can be selected to achieve the purpose of this embodiment according to actual needs. Those skilled in the art can understand and implement this without any creative effort.

[0075] Through the above description of the embodiments, those skilled in the art can clearly understand that each embodiment can be implemented by means of software plus necessary general-purpose hardware platforms, and of course, it can also be implemented by hardware. Based on this understanding, the above technical solutions, in essence or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product can be stored in a computer-readable storage medium, such as ROM / RAM, magnetic disk, optical disk, etc., and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute the methods described in the various embodiments or some parts of the embodiments.

[0076] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A finite-time sliding mode control method for integrated attitude and orbit control of an aircraft, characterized in that, This includes the following integrated attitude and orbit control process for high-speed maneuvering aircraft: A relative attitude and orbit integrated kinematic and dynamic model of the aircraft and the target is established based on the Lie group SE(3); A non-singular terminal sliding surface is defined, and a hyperbolic function is used to replace the sign function to construct the sliding surface expression; the sliding surface is constructed using the following formula: in, , These are the weighting coefficients. The coefficients are non-linear exponential coefficients. For speed tracking error, For exponential coordinates based on the Lie group SE(3); Differentiating with respect to the sliding surface yields Where sech() represents the hyperbolic secant function, Indicates the defined part of the system. For system interference, It is the inertial tensor matrix composed of the aircraft's mass and moment of inertia. For system control input, This is the velocity transformation matrix; A finite-time sliding mode controller is configured in conjunction with an interference observer, wherein the interference observer is used to estimate system interference and generate interference estimates. The control input is generated based on the sliding mode controller to achieve convergence of attitude and trajectory errors within a finite time.

2. The finite-time sliding mode control method for integrated attitude and orbit control of an aircraft according to claim 1, characterized in that: The setting of a non-singular terminal sliding surface includes constructing a sliding surface based on generalized velocity error and using a hyperbolic secant function to suppress chattering during the sliding mode control process.

3. The finite-time sliding mode control method for integrated attitude and orbit control of an aircraft according to claim 1, characterized in that: The interference observer generates an estimate of the system interference and its derivative through multi-level correction terms, adjusts the observer output using nonlinear gain coefficients and adaptive parameter matrices, and then constructs the corresponding observer error equation: in, They are respectively The estimated value, , , for The corresponding time derivative, This is a first-level correction item. This is a second-level correction item. , , For nonlinear gain coefficients, For adaptive parameter matrix, It is a symbolic function.

4. The finite-time sliding mode control method for integrated attitude and orbit control of an aircraft according to claim 1, characterized in that: The expression for the finite-time sliding mode controller includes a sliding surface term, a disturbance estimation compensation term for the disturbance observer, control coefficients, the minimum eigenvalue function of the system inertia tensor matrix, and a hyperbolic function term for smoothing the control input, as shown in the following equation. In the formula, , , For control coefficients, It is the minimum eigenvalue function.

5. The finite-time sliding mode control method for integrated attitude and orbit control of an aircraft according to claim 1, characterized in that: The finite-time convergence is proven by Lyapunov stability theory, and the convergence time is determined by the initial state, sliding surface parameters, and control coefficients.

6. An integrated attitude and orbit control system for an aircraft, characterized in that, Includes the following modules: The kinematics and dynamics modeling module is used to construct an integrated attitude and orbit model of the aircraft and the target based on the Lie group SE(3); The interference observer module is used to estimate and output interference estimates; A non-singular terminal sliding mode controller module is used to construct a sliding mode surface using a hyperbolic function and generate control inputs based on the disturbance estimate. The stability analysis module is used to verify the finite-time convergence of the system state.

7. The finite-time sliding mode control method for integrated attitude and orbit control of an aircraft according to claim 1, characterized in that: The non-singular terminal sliding mode controller module is configured to embed the hyperbolic secant function into the sliding surface derivative function to suppress chattering, adjust the estimation accuracy of the interference observer through a multi-level correction mechanism, and ensure finite-time convergence by combining control coefficients and eigenvalue functions.

8. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the program, it implements the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft as described in any one of claims 1 to 5.

9. A non-transitory computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by the processor, it implements the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft as described in any one of claims 1 to 5.

10. A computer program product, comprising a computer program, characterized in that: When the computer program is executed by the processor, it implements the finite-time sliding mode control method for integrated attitude and orbit control of an aircraft as described in any one of claims 1 to 5.

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