Predefined time contact racemization control method of flexible fault satellite and related device

Abstract

The invention discloses a predefined time contact racemization control method for a flexible fault satellite and a related device, and the method comprises the steps: obtaining the positions and attitudes of a service spacecraft and the flexible fault satellite, and building a dynamic model of the service spacecraft and the flexible fault satellite; designing predefined time based on adaptive compensation according to the state of the service spacecraft in combination with the interference estimation value; the method comprises the steps of establishing a dynamic model of a flexible sailboard of a flexible fault satellite, establishing a dynamic model of a rigid-flexible coupling system of the flexible fault satellite by adopting a Lagrange multiplier method according to predefined time and the state of the flexible fault satellite, and performing contact racemization control according to the state of the rigid-flexible coupling system of the flexible fault satellite. According to the method, the tedious parameter adjustment process is avoided, the stability of the service spacecraft can be realized within fixed time according to the actual racemization task, the calculation complexity is reduced, and the requirement of the real-time racemization task is easily met.

Description

Technical Field

[0001] The invention belongs to the technical field of spacecraft control, and relates to a predefined time contact derotation control method for a flexible fault satellite and a related device. Background Art

[0002] In the past half century, humans have launched more than 15,000 artificial satellites, of which more than 80% have failed due to fuel exhaustion and mechanical failure. In addition to failed satellites, there are still a large number of space debris such as rocket upper stages, spacecraft disintegration and collision derivatives scattered in space. These space debris have occupied precious orbital resources for a long time, not only threatening the normal operation of on-orbit spacecraft, but also posing safety risks to future space activities. Therefore, the active removal technology of garbage has become a key issue in on-orbit services. Among the many active removal technologies, the use of space robots to capture failed satellites has attracted widespread attention from scholars. However, failed satellites usually carry large-scale flexible sails, which are affected by the space environment and their own residual angular momentum, and exhibit complex free tumbling motion, which poses a huge challenge to direct capture missions. Therefore, it is a prerequisite for safe capture to implement a derotation operation on the satellite to attenuate its angular velocity to a range that meets the capture requirements. In the past few decades, scholars have proposed many contact and non-contact derotation methods, among which the contact derotation method includes electromagnetic eddy current method, thruster plume method and laser ablation method, etc. However, these methods provide relatively small control force and it is difficult to achieve an efficient derotation process. Compared with the non-contact derotation method, the contact derotation method can provide greater control force, especially for large flexible satellites, it can achieve higher derotation efficiency. Installing a flexible operating rod on a service spacecraft as an end effector, pushing into the satellite and contacting the edge of its sailboard to perform the derotation operation can not only avoid the risk of rigid collision, but also has the characteristics of simple device and flexible operation. It is an efficient and safe method.

[0003] However, the large flexible deformation of the flexible operating rod causes the dynamics of the service spacecraft to present complex nonlinear and strong coupling characteristics, and the strong disturbance caused by contact easily leads to the instability of the service spacecraft. Therefore, the design of the derotation controller becomes extremely complex and challenging.

[0004] At present, most deceleration control methods can only ensure asymptotic convergence, which is difficult to meet the deceleration tasks with strict time constraints, and there are still certain limitations. See the literature: Dai H, Chen H, Yue X. Dynamic analysis of detumbling a rotating satellite using flexible deceleration rod[J]. NonlinearDynamics, 2022, 108(4): 3331-3345. Therefore, for the deceleration process of flexible satellites, a finite time convergence control strategy must be designed, see the literature: Wang Hongwei, Dai Honghua, Chen Hao, Yue Xiaokui. Contact deceleration output feedback control of space tumbling targets under coupled disturbances[J]. Journal of Astronautics, 2023, 44(10): 1564-1574. However, the convergence time of this method is seriously dependent on the initial state, which may be infinitely extended as the initial state increases, and the upper bound is determined by multiple control parameters, which increases the difficulty of selecting control parameters, and is therefore not suitable for actual deceleration tasks. The predefined time control method can solve this problem, but most of the existing methods are only for rigid spacecraft, ignoring the coupling effect of flexible operating rods. Therefore, it is particularly urgent to design a predefined time control strategy for serving spacecraft during the derotation process. Summary of the invention

[0005] The purpose of the present invention is to provide a predefined time contact derotation control method and related devices for a flexible fault satellite, so as to solve the problem that the prior art only targets rigid spacecraft and ignores the coupling effect of the flexible operating rod.

[0006] In order to achieve the above object, the present invention adopts the following technical solutions: Predefined time contact derotation control method for flexible fault satellite, including: Obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish the dynamic model of the service spacecraft and the flexible fault satellite; According to the status of the service spacecraft and the interference estimation value, a predefined time based on adaptive compensation is designed; The dynamic model of the flexible sailboard of the flexible fault satellite is established. According to the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, contact derotation control is performed.

[0007] Furthermore, the dynamic model of the service spacecraft is:

[0008]

[0009] in, represents the mass of the service spacecraft base, represents the position vector from the center of mass of the service spacecraft to the origin of the inertial system, represents the service spacecraft moment of inertia, represents the Euler angle vector of the service spacecraft, represents the attitude angular velocity vector of the service spacecraft, represents the control torque acting on the service spacecraft base, Control forces acting on the service vehicle matrix, matrix Expressed as .

[0010] Furthermore, the dynamic model of the flexible fault satellite is:

[0011] in, represents the mass of the soft failure satellite, is the position vector of the flexible fault satellite, represents the Euler angle vector of the soft fault satellite, represents the angular velocity vector of the flexible fault satellite, express The antisymmetric matrix of Expressed as .

[0012] Furthermore, the interference estimation value is:

[0013] in, is the observer gain, Indicates the status of the anti-windup compensator.

[0014] Furthermore, the dynamic model of the flexible sailboard of the flexible fault satellite is:

[0015] in, is the unit mass matrix, is the system generalized coordinate, The generalized external force matrix, Damping force matrix, is the elastic force matrix.

[0016] Furthermore, the process of establishing the dynamic model of the flexible sailboard of the flexible fault satellite is as follows: The flexible sailboard of the flexible fault satellite is described by a rectangular four-node unit, and the kinetic energy of the flexible sailboard unit is calculated by using the displacement field function of any point in the flexible sailboard unit in the global coordinate system. According to Kirchhoff plate theory, the strain energy of the flexible sailboard is calculated, wherein the strain energy of the flexible sailboard includes the shear strain in the longitudinal direction of the mid-surface and the bending and torsion of the mid-surface; According to the kinetic energy and strain energy of the flexible sailboard unit, the dynamic model of the flexible sailboard is derived using the Euler-Lagrange equation.

[0017] Furthermore, the dynamic equation of the rigid-flexible coupling system of the flexible fault satellite is:

[0018] in, Indicates the system quality, represents the system generalized coordinates, represents the Lagrange multiplier vector, represents the constraint equation, represents the generalized force vector.

[0019] Predefined time contact de-rotation control system for fail-proof satellites, including: A modeling module, wherein the modeling module is used to obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish a dynamic model of the service spacecraft and the flexible fault satellite; A design module, the design module is used to design a predefined time based on adaptive compensation according to the state of the service spacecraft and the interference estimation value; A control module is used to establish a dynamic model of the flexible sailboard of the flexible fault satellite, establish a dynamic model of the rigid-flexible coupling system of the flexible fault satellite using the Lagrange multiplier method according to the predefined time and the state of the flexible fault satellite, and perform contact derotation control according to the state of the rigid-flexible coupling system of the flexible fault satellite.

[0020] A terminal device comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the steps of the method are implemented when the processor executes the computer program.

[0021] A computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the method are implemented.

[0022] Compared with the prior art, the present invention has the following beneficial effects: The present invention provides a predefined time contact derotation control method for a flexible fault satellite. By acquiring the position and attitude of the service spacecraft and the flexible fault satellite, a dynamic model of the service spacecraft and the flexible fault satellite is established. According to the state of the service spacecraft, combined with the interference estimation value, a predefined time based on adaptive compensation is designed. By constructing a new disturbance observer and an adaptive compensation framework, while reducing the complexity of the control algorithm, the anti-interference ability of the system is greatly improved and the fine control performance is guaranteed, providing a solution with both high real-time performance and strong robustness for tasks such as derotation of non-cooperative targets in space, and realizing the robustness of the service spacecraft under strong disturbance. Then, a dynamic model of the flexible sailboard of the flexible fault satellite is established, and according to the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, contact derotation control is performed, which ensures that the service spacecraft achieves the convergence of the predefined time, and ensures that the derotation of the flexible fault satellite is realized efficiently and effectively. The present invention has a simple structure, reduces computational complexity, and only needs to adjust one parameter to achieve fixed-time convergence, thus avoiding tedious parameter adjustment processes. Therefore, the stability of the service spacecraft can be achieved within a fixed time according to the actual derotation task, and it is easy to meet the requirements of real-time derotation tasks. The robust control method of the predefined time of the present invention introduces a time-varying gain mechanism, breaks through the theoretical limitations of traditional fixed-time control, and realizes engineering-oriented precise control of preset time. The upper limit of the system convergence time can be directly set according to task requirements, and the preset time stable control of the service spacecraft and the high-precision derotation operation of the flexible fault satellite can be realized under strong time-varying interference and uncertainty conditions, which significantly improves the engineering applicability of on-orbit operations and has strong engineering application value. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings required for use in the embodiments are briefly introduced below. It should be understood that the following drawings only show certain embodiments of the present invention and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other related drawings can be obtained based on these drawings without creative work.

[0024] Figure 1 This is a flow chart of the predefined time contact derotation control method for a flexible fault satellite of the present invention.

[0025] Figure 2 It is a structural schematic diagram of the contact derotation process of the flexible fault satellite of the present invention.

[0026] Figure 3 This is a diagram of the attitude change history of the service spacecraft of the present invention.

[0027] Figure 4This is a diagram of the thrust change history of the service spacecraft of the present invention.

[0028] Figure 5 This is a diagram of the angular velocity change history of the service spacecraft of the present invention.

[0029] Figure 6 This is a diagram showing the angular velocity change history of a flexible fault satellite according to the present invention.

[0030] Figure 7 This is a diagram showing the position change history of a flexible fault satellite according to the present invention.

[0031] Figure 8 This is a translational velocity history diagram of a flexible fault satellite according to the present invention.

[0032] Fig. 9 This is a schematic diagram of the structure of a predefined time contact derotation control system for a flexible fault satellite according to a preferred embodiment of the present invention.

[0033] Fig.10 It is a schematic diagram of the structure of an electronic device according to a preferred embodiment of the present invention. DETAILED DESCRIPTION

[0034] The following is a description of exemplary embodiments of the present application in conjunction with the accompanying drawings, including various details of the embodiments of the present application to facilitate understanding, which should be considered as merely exemplary. Therefore, it should be recognized by those of ordinary skill in the art that various changes and modifications can be made to the embodiments described herein without departing from the scope and spirit of the present application. Similarly, for the sake of clarity and conciseness, the description of well-known functions and structures is omitted in the following description.

[0035] Obviously, the described embodiments are only part of the embodiments of the present application, but not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without creative work are within the scope of protection of the present application.

[0036] It should be noted that the terminals involved in the embodiments of the present application may include but are not limited to mobile phones, personal digital assistants (PDAs), wireless handheld devices, tablet computers, personal computers (PCs), MP3 players, MP4 players, wearable devices (for example, smart glasses, smart watches, smart bracelets, etc.), smart home devices and other smart devices.

[0037] In addition, the term "and / or" in this article is only a description of the association relationship between the associated objects, indicating that there can be three relationships. For example, A and / or B can represent: A exists alone, A and B exist at the same time, and B exists alone. In addition, the character " / " in this article generally indicates that the associated objects before and after are in an "or" relationship.

[0038] The present invention is further described in detail below in conjunction with the accompanying drawings: See also Figure 1 The present invention provides a predefined time contact derotation control method for a flexible fault satellite, which specifically includes the following steps: Step 1: Obtain the position and attitude of the service spacecraft and establish a dynamic model of the service spacecraft.

[0039] The dynamic equations of position and attitude of the service spacecraft are expressed as (1) (2) Where: represents the mass of the service spacecraft base, represents the position vector from the center of mass of the service spacecraft to the origin of the inertial system, represents the service spacecraft moment of inertia, represents the Euler angle vector of the service spacecraft, represents the attitude angular velocity vector of the service spacecraft, represents the control torque acting on the service spacecraft base, Control forces acting on the service vehicle matrix, matrix Expressed as (3) In order to facilitate the establishment of the racemization control system, the dynamic equation is written as the following Euler-Langerange system, expressed as (4) Where: , .

[0040] Step 2: According to the status of the service spacecraft and the interference estimation value, a predefined time control based on adaptive compensation is designed.

[0041] Redefine the service spacecraft status as and , the service spacecraft dynamics system can be written in the following state space form (5) Where: , , represents the coupled disturbance acting on the spacecraft system, where and is the Jacobian matrix, which represents the constraint equations on the attitude variables and position variables The partial derivative of .

[0042] Tracking error is defined as (6) In the formula, and is the expected trajectory and its first-order derivative.

[0043] Step 1: The backstepping variable is defined as (7) Where: is a virtual control input, designed as (8) In the formula, , , , Indicates the preset time.

[0044] The first-order derivative with respect to time is (9) Where: , , , It is the status of the anti-saturation compensator.

[0045] Define the first Lyapunov function (10) Taking the first-order derivative of equation (9) with respect to time, we can obtain (11) Substituting the virtual control input into equation (10) yields (12) Where: Is a normal number.

[0046] Step 2: Define the second Lyapunov function (13) Taking the first-order derivative of the second equation of formula (7) with respect to time, we can obtain (14) Taking the derivative of equation (13) with respect to time, we can obtain (15) The actual control input can be designed as (16) (17) Where: is the normal number to be designed.

[0047] Interference estimation algorithm design (18) Where: is the observer gain.

[0048] The estimation algorithm is designed as (19) Where: express Estimation error.

[0049] The interference estimate is designed to be (20) The error dynamics is (twenty one) Substituting formula (15) into formula (16), we can get (twenty two) Where: and Is a normal number.

[0050] Step 3: System stability analysis.

[0051] Consider the following Lyapunov function (twenty three) Taking the first-order derivative of equation (23) with respect to time, we can obtain (twenty four) Where: .

[0052] It can be concluded that: , and At preset time Converges to a smaller region of equilibrium points. Therefore, the service spacecraft state tracks the desired state within a preset time, and the designed disturbance observer estimates the coupled disturbance at a preset time.

[0053] Step 4: Establish the dynamic model of the flexible fault satellite and the flexible sailboard. According to the predefined time and the state of the flexible fault satellite, use the Lagrange multiplier method to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, perform contact derotation control.

[0054] A flexible sailboard dynamic model based on the absolute node coordinate method is established by using a rectangular four-node unit to describe the flexible board. The unit contains four nodes, and the coordinate vector of each node has 9 components. The node coordinate vector of the unit contains 36 components in total, so the unit has 36 degrees of freedom. In the global coordinate system, the displacement field function can be expressed as (25) Where: is the shape function, where , , , , , , , , , , , , , is the unit length, is the unit width, are the nodal coordinates of the element.

[0055] Assume that any node coordinate array in the flexible sailboard unit is expressed as (26) Where: is the global position vector of point A, for For local coordinates The bias leads to for For local coordinates The partial derivative of is the global position vector of point B, for For local coordinates The bias leads to for For local coordinates The partial derivative of is the global position vector of point C, for For local coordinates The bias leads to for For local coordinates The partial derivative of is the global position vector of point D, for For local coordinates The bias leads to for For local coordinates The partial derivative of .

[0056] but Expressed as (27) Where: , , , are the node coordinates of unit nodes A, B, C, and D respectively.

[0057] The unit kinetic energy is expressed as (28) Where: represents the element mass matrix, is the cell density, is the unit thickness.

[0058] According to Kirchhoff plate theory, the strain energy of a sailboard is generated by two parts: one is generated by the longitudinal shear strain of the mid-surface, and the other is generated by the bending and torsion of the mid-surface.

[0059] The strain vector of the mid-surface is expressed as (29) Where: , , , , .

[0060] The curvature vector of the mid-surface is expressed as (30) Where: , , , , , , , .

[0061] The strain energy of the sailboard element is expressed as (31) Where: , , , , , , , , , , , , , , . and is the elastic coefficient matrix, expressed as (32) Where: is the elastic modulus, is Poisson's ratio.

[0062] The elastic force calculation is expressed as (33) Considering the viscoelasticity of the material, the element shear strain and the generalized forces generated by bending and torsion are expressed as (34) Where: and is the corresponding damping coefficient.

[0063] Using the Euler-Lagrange equation, the dynamic equation of the windsurfing system can be derived as follows: (35) in, is the unit mass matrix, is the system generalized coordinate, The generalized external force matrix, Damping force matrix, is the elastic force matrix.

[0064] The satellite attitude dynamics equation is expressed as (36) Where: Indicates quality, is the position vector, represents the Euler angle vector, represents the angular velocity vector, express The antisymmetric matrix of Expressed as (37) The Lagrange multiplier method is used to establish the dynamic equation of the flexible satellite rigid-flexible coupling system, which is expressed as a differential-algebraic equation group of index-3: (38) Where: For system quality, is the system generalized coordinate, is the Lagrange multiplier vector, is the constraint equation, represents the generalized force vector.

[0065] In order to facilitate the solution of the above equations, they are usually converted into ordinary differential equations of index-1, and the Baumgarte stable method is used to solve the equation system: (39) Where: , , , and Feedback control parameters to limit speed and position constraint violations.

[0066] The present invention is further described in detail below through specific embodiments: Embodiment 1: In this embodiment, the predefined time contact derotation control method of the flexible fault satellite specifically includes the following steps: Step 1: Obtain the position and attitude of the service spacecraft and establish a dynamic model of the service spacecraft.

[0067] The dynamic equations of position and attitude of the service spacecraft are expressed as (1) Where: Indicates the mass of the base, represents the position vector from the center of mass of the service spacecraft to the origin of the inertial system, is the spacecraft moment of inertia, represents the Euler angle vector, which is used to describe the attitude of the spacecraft. represents the spacecraft's attitude angular velocity vector, represents the control torque acting on the matrix, and the matrix Expressed as (2) The attitude dynamics equation is written as the following Euler-Langerange equation form, and the service spacecraft attitude dynamics model is expressed as (3) Where: , , represents the position vector from the center of mass of the service spacecraft to the origin of the inertial system, Controlling forces acting on the substrate.

[0068] Step 2: According to the status of the service spacecraft and the interference estimation value, a predefined time control based on adaptive compensation is designed.

[0069] Redefine the service spacecraft status as and , the service spacecraft dynamics system can be written in the following state space form (4) Where: , , represents the coupled disturbance acting on the spacecraft system, where and is the Jacobian matrix, which represents the constraint equations on the attitude variables and position variables The partial derivative of .

[0070] (5) Among them, the virtual control input Designed for (6) Where: , , , Indicates the preset time.

[0071] The actual control input can be designed as (7) (8) Where: is the normal number to be designed.

[0072] Interference estimation algorithm design (9) Where: is the observer gain.

[0073] The estimation algorithm is designed as (10) Where: express Estimation error.

[0074] The interference estimate is designed to be (11) Step 3: Establish the dynamic model of the flexible fault satellite and the flexible sailboard. According to the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, contact derotation control is performed, such as Figure 2 shown.

[0075] A flexible sailboard is described by a rectangular four-node unit, and a flexible sailboard dynamic model based on the absolute node coordinate method is established. The unit contains four nodes, and the coordinate vector of each node has 9 components. The node coordinate vector of the unit contains 36 components in total, so the unit has 36 degrees of freedom. In the global coordinate system, the displacement field function can be expressed as (12) Where: is the shape function, where , , , , , , , , , , , , , is the unit length, is the unit width, are the nodal coordinates of the element.

[0076] Assume that any node coordinate array in the sailboard unit Expressed as (13) Where: , , , are the node coordinates of unit nodes A, B, C, and D respectively.

[0077] The unit kinetic energy is expressed as (14) Where: represents the element mass matrix, is the cell density, is the unit thickness.

[0078] According to Kirchhoff plate theory, the strain energy of a sailboard is generated by two parts: one is generated by the longitudinal shear strain of the mid-surface, and the other is generated by the bending and torsion of the mid-surface.

[0079] The curvature vector of the mid-surface is expressed as (15) Where: , , , , , , , .

[0080] The strain energy of the sailboard element is expressed as (16) Where: , , , , , , , , , , , , , , . and is the elastic coefficient matrix, expressed as (17) Where: is the elastic modulus; is Poisson's ratio.

[0081] The elastic force calculation is expressed as (18) Considering the viscoelasticity of the material, the element shear strain and the generalized forces generated by bending and torsion are expressed as (19) Where: and is the corresponding damping coefficient.

[0082] Using the Euler-Lagrange equation, the dynamic equation of the windsurfing system can be derived as follows: (20) in, is the unit mass matrix, is the system generalized coordinate, The generalized external force matrix, Damping force matrix, is the elastic force matrix.

[0083] The satellite attitude dynamics equation is expressed as (twenty one) Where: Indicates quality, is the position vector, represents the Euler angle vector, represents the angular velocity vector, express The antisymmetric matrix of Expressed as (twenty two) The Lagrange multiplier method is used to establish the dynamic equation of the flexible satellite rigid-flexible coupling system, which is expressed as a differential-algebraic equation group of index-3: (twenty three) Where: For system quality, is the system generalized coordinate, is the Lagrange multiplier vector, is the constraint equation, represents the generalized force vector.

[0084] In order to facilitate the solution of the above equations, they are usually converted into ordinary differential equations of index-1, and the Baumgarte stable method is used to solve the equation system: (twenty four) Where: , , , and Feedback control parameters to limit speed and position constraint violations.

[0085] Step 4: Perform numerical simulation.

[0086] The parameters of the service spacecraft dynamics model are selected as follows: moment of inertia ,quality . The initial value is set to: ; ; ; .

[0087] The dynamic simulation parameters of the flexible fault satellite are shown in Table 1 and Table 2: Table 1 Physical parameters of flexible satellite sailboard

[0088] Table 2 Physical parameters of the rigid body of the flexible satellite center

[0089] The controller parameters are selected as follows: ; ; ; ; .

[0090] The numerical simulation results are as follows Figures 3 to 8 shown. Figure 3 The change process diagram of the service spacecraft attitude angle is given. It can be seen that the service spacecraft has undergone 9 derotation processes, and the attitude angle can converge to the equilibrium position after each derotation. Figure 4 The change history of the thruster is given, and it can be seen that in each de-rotation process, the thrust is limited to the set saturation range. Figure 5 The graph depicts the change history of the angular velocity of the service spacecraft. It can be seen that the angular velocity can quickly converge to the equilibrium position to maintain the desired orientation of the operating rod.

[0091] Figure 6 The change history of the satellite angular velocity is given. It can be seen that the satellite angular velocity is reduced to a range that satisfies subsequent capture after 9 de-rotation processes. Figure 7 The change history of the satellite position is plotted, and it can be seen that the satellite X and X direction positions gradually move away from each other, which is due to the contact force. Figure 8 The change history of the satellite's translational velocity is given. It can be seen that the translational velocity presents an increasing trend and has an oscillating phenomenon.

[0092] The present invention also provides a predefined time contact derotation control system for a flexible fault satellite, such as Fig. 9 As shown, the system includes: a modeling module, a design module and a control module.

[0093] A modeling module, wherein the modeling module is used to obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish a dynamic model of the service spacecraft and the flexible fault satellite; A design module, the design module is used to design a predefined time based on adaptive compensation according to the state of the service spacecraft and the interference estimation value; A control module is used to establish a dynamic model of the flexible sailboard of the flexible fault satellite, establish a dynamic model of the rigid-flexible coupling system of the flexible fault satellite using the Lagrange multiplier method according to the predefined time and the state of the flexible fault satellite, and perform contact derotation control according to the state of the rigid-flexible coupling system of the flexible fault satellite.

[0094] It can be understood that the predefined time contact derotation control system for a flexible fault satellite provided by the present invention corresponds to the predefined time contact derotation control method for a flexible fault satellite provided in the aforementioned embodiments. The relevant technical features of the predefined time contact derotation control system for a flexible fault satellite can refer to the relevant technical features of the predefined time contact derotation control method for a flexible fault satellite, which will not be repeated here.

[0095] Another object of the present invention is to provide an electronic device, such as Fig.10 As shown, it includes a memory, a processor, and a computer program stored in the memory and executable on the processor, and the processor executes the steps of the predefined time contact derotation control method of the flexible failure satellite.

[0096] The predefined time contact derotation control method of the flexible fault satellite comprises the following steps: Obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish the dynamic model of the service spacecraft and the flexible fault satellite; According to the status of the service spacecraft and the interference estimation value, a predefined time based on adaptive compensation is designed; The dynamic model of the flexible sailboard of the flexible fault satellite is established. According to the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, contact derotation control is performed.

[0097] A fourth object of the present invention is to provide a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the steps of the predefined time contact derotation control method for the flexible fault satellite.

[0098] The predefined time contact derotation control method of the flexible fault satellite comprises the following steps: Obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish the dynamic model of the service spacecraft and the flexible fault satellite; According to the status of the service spacecraft and the interference estimation value, a predefined time based on adaptive compensation is designed; The dynamic model of the flexible sailboard of the flexible fault satellite is established. According to the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, contact derotation control is performed.

[0099] It should be understood by those skilled in the art that the embodiments of the present invention may be provided as methods, systems or computer program products. Therefore, the present invention may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware aspects. Moreover, the present invention may take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0100] The present invention is described with reference to flowcharts and / or block diagrams of methods, devices (systems), and computer program products according to embodiments of the present invention. It should be understood that each process and / or block in the flowchart and / or block diagram, as well as the combination of processes and / or blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0101] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0102] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0103] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the above embodiments, ordinary technicians in the relevant field should understand that the specific implementation methods of the present invention can still be modified or replaced by equivalents. Any modification or equivalent replacement that does not depart from the spirit and scope of the present invention should be covered within the scope of protection of the claims of the present invention.

Claims

1. A predefined time contact derotation control method for a flexible fault satellite, characterized in that: include: Obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish the dynamic model of the service spacecraft and the flexible fault satellite; According to the status of the service spacecraft and the interference estimation value, a predefined time based on adaptive compensation is designed; The dynamic model of the flexible sailboard of the flexible fault satellite is established. According to the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish the dynamic model of the rigid-flexible coupling system of the flexible fault satellite. According to the state of the rigid-flexible coupling system of the flexible fault satellite, contact derotation control is performed.

2. The predefined time contact derotation control method for a flexible fault satellite according to claim 1, characterized in that: The dynamic model of the service spacecraft is: in, represents the mass of the service spacecraft base, represents the position vector from the center of mass of the service spacecraft to the origin of the inertial system, is the service spacecraft moment of inertia, represents the Euler angle vector of the service spacecraft, represents the attitude angular velocity vector of the service spacecraft, represents the control torque acting on the service spacecraft base, Control forces acting on the service vehicle matrix, matrix Expressed as .

3. The predefined time contact derotation control method for a flexible fault satellite according to claim 1, characterized in that: The dynamic model of the flexible fault satellite is: in, represents the mass of the soft failure satellite, is the position vector of the flexible fault satellite, represents the Euler angle vector of the soft fault satellite, represents the angular velocity vector of the flexible fault satellite, express The antisymmetric matrix of Expressed as .

4. The predefined time contact derotation control method for a flexible fault satellite according to claim 1, characterized in that: The interference estimate is: in, is the observer gain, Indicates the status of the anti-windup compensator.

5. The predefined time contact derotation control method for a flexible fault satellite according to claim 1, characterized in that: The dynamic model of the flexible sailboard of the flexible fault satellite is: in, is the unit mass matrix, is the system generalized coordinate, The generalized external force matrix, Damping force matrix, is the elastic force matrix.

6. The predefined time contact derotation control method for a flexible fault satellite according to claim 1, characterized in that: The process of establishing the dynamic model of the flexible sailboard of the flexible fault satellite is as follows: The flexible sailboard of the flexible fault satellite is described by a rectangular four-node unit, and the kinetic energy of the flexible sailboard unit is calculated by using the displacement field function of any point in the flexible sailboard unit in the global coordinate system. According to Kirchhoff plate theory, the strain energy of the flexible sailboard is calculated, wherein the strain energy of the flexible sailboard includes the shear strain in the longitudinal direction of the mid-surface and the bending and torsion of the mid-surface; According to the kinetic energy and strain energy of the flexible sailboard unit, the dynamic model of the flexible sailboard is derived using the Euler-Lagrange equation.

7. The predefined time contact derotation control method for a flexible fault satellite according to claim 1, characterized in that: The dynamic equation of the rigid-flexible coupling system of the flexible fault satellite is: in, Indicates the system quality, represents the system generalized coordinates, represents the Lagrange multiplier vector, represents the constraint equation, represents the generalized force vector.

8. Predefined time contact derotation control system for flexible failure satellites, characterized in that, include: A modeling module, wherein the modeling module is used to obtain the position and attitude of the service spacecraft and the flexible fault satellite, and establish a dynamic model of the service spacecraft and the flexible fault satellite; A design module, the design module is used to design a predefined time based on adaptive compensation according to the state of the service spacecraft and the interference estimation value; A control module is used to establish a dynamic model of the flexible sailboard of the flexible fault satellite, establish a dynamic model of the rigid-flexible coupling system of the flexible fault satellite using the Lagrange multiplier method according to the predefined time and the state of the flexible fault satellite, and perform contact derotation control according to the state of the rigid-flexible coupling system of the flexible fault satellite.

9. A terminal 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 computer program, the steps of the method according to any one of claims 1 to 7 are implemented.

10. A computer-readable storage medium storing a computer program, characterized in that: When the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 7 are implemented.

Images