Predefined time contact despin control method for flexible failed satellite and related device
By establishing a dynamic model and adaptive compensation framework for flexible fault satellites and designing a predefined time control method, the problem of convergence time dependence on initial state in the existing technology of flexible satellite despin control is solved, and efficient and robust despin control under strong disturbances is realized.
Patent Information
- Application Number
- CN202510441152.3
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2025-04-09
- Publication Date
- 2026-05-15
- Estimated Expiration
- 2045-04-09
AI Technical Summary
Existing technologies make it difficult to design predefined time-contact despin control methods for flexible fault satellites, neglecting the coupling effect of flexible control levers, resulting in the control strategy's convergence time depending on the initial state and making it difficult to meet strict time constraints.
By establishing dynamic models of service spacecraft and flexible fault satellites, a predefined time control based on adaptive compensation is designed. A dynamic model of the rigid-flexible coupled system of the flexible fault satellite is established by combining the Lagrange multiplier method. A novel disturbance observer and adaptive compensation framework are used to achieve contact despin control.
It achieves robustness and precise control performance for spacecraft under strong disturbances, ensures efficient despinning of flexible fault satellites within a fixed time, reduces the complexity of control algorithms, and is suitable for practical despinning missions.
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Figure CN120096835B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of spacecraft control technology and relates to a predefined time contact despin control method and related devices for flexible fault satellites. Background Technology
[0002] Over the past half-century, humanity has launched more than 15,000 artificial satellites, of which over 80% have failed due to fuel exhaustion and mechanical malfunctions. Besides failed satellites, space debris, including rocket upper stages, spacecraft disintegrations, and collision byproducts, remains scattered throughout the atmosphere. This space debris occupies valuable orbital resources, threatening not only the normal operation of spacecraft in orbit but also posing safety risks to future space activities. Therefore, active debris removal technology has become a critical issue in on-orbit servicing. Among the many active debris removal technologies, the use of space robots to capture failed satellites has attracted widespread attention. However, failed satellites typically carry large flexible solar panels, which, influenced by the space environment and their own residual angular momentum, exhibit complex free tumbling motions, posing a significant challenge to direct capture missions. Therefore, performing despinning operations to reduce the satellite's angular velocity to a range suitable for capture is a prerequisite for safe capture. Over the past few decades, scholars have proposed numerous contact and non-contact despinning methods. Contact despinning methods include electromagnetic eddy current methods, thruster plume methods, and laser ablation methods. However, these methods provide relatively small control forces, making it difficult to achieve efficient despinning processes. Compared to non-contact despinning methods, contact despinning methods can provide greater control forces, especially for large flexible satellites, achieving higher despinning efficiency. Installing a flexible control lever on the servicing spacecraft as an end effector, and then approaching the satellite and contacting the edge of its solar panels to perform despinning operations, not only avoids the risk of rigid collisions but also features simple equipment and flexible operation, making it an efficient and safe method.
[0003] However, the large deformation of the flexible control lever causes the servicing spacecraft's dynamics to exhibit complex nonlinear and strongly coupled characteristics. Furthermore, the strong disturbances caused by contact can easily lead to instability in the servicing spacecraft. Therefore, designing a despin controller becomes extremely complex and challenging.
[0004] Currently, most despin control methods can only ensure asymptotic convergence, which is difficult to meet the despin tasks with strict time constraints and still has certain limitations. See reference: 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 despin process of flexible satellites, a finite-time convergence control strategy must be designed, see reference: Wang Hongwei, Dai Honghua, Chen Hao, Yue Xiaokui. Contact despin 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 depends heavily on the initial state and may extend infinitely with the increase of the initial state, and the upper bound is determined by multiple control parameters, which increases the difficulty of selecting control parameters and is therefore not suitable for practical despin tasks. Predefined timing control methods can solve this problem; however, most existing methods only apply to rigid spacecraft, neglecting the coupling effect of flexible control levers. Therefore, designing a predefined timing control strategy to serve the spacecraft during despinning is particularly urgent. Summary of the Invention
[0005] The purpose of this invention is to provide a predefined time contact despin control method and related device for flexible fault satellites, which solves the problem that the existing technology only targets rigid spacecraft and ignores the coupling effect of flexible control levers.
[0006] To achieve the above objectives, the present invention employs the following technical solution:
[0007] A predefined time-contact despin control method for flexible fault satellites includes:
[0008] Obtain the position and attitude of the servicing spacecraft and the flexible fault satellite, and establish dynamic models of the servicing spacecraft and the flexible fault satellite;
[0009] Based on the state of the servicing spacecraft and the interference estimate, a predefined time based on adaptive compensation is designed.
[0010] A dynamic model of the flexible solar panel of the flexible fault satellite is established. Based on the predefined time and the state of the flexible fault satellite, a dynamic model of the rigid-flexible coupled system of the flexible fault satellite is established using the Lagrange multiplier method. Contact despin control is then performed based on the state of the rigid-flexible coupled system of the flexible fault satellite.
[0011] Furthermore, the dynamic model of the service spacecraft is as follows:
[0012]
[0013]
[0014] in, Indicates the mass of the spacecraft base. This represents the position vector from the center of mass of the serving spacecraft to the origin of the inertial frame. Indicates the moment of inertia of the service spacecraft. Represents the Euler angle vector of the servicing spacecraft. This represents the attitude angular velocity vector of the servicing spacecraft. This represents the control torque acting on the service spacecraft matrix. Control forces acting on the service spacecraft matrix, matrix Represented as .
[0015] Furthermore, the dynamic model of the flexible fault satellite is as follows:
[0016]
[0017] in, Indicates the mass of a satellite with a flexible fault. For the position vector of the flexible fault satellite, The Euler angle vector representing a satellite with a flexible fault. The vector representing the angular velocity of a satellite with a flexible fault. express antisymmetric matrix, Represented as .
[0018] Furthermore, the interference estimate is:
[0019]
[0020] in, For observer gain, This indicates the state of the anti-saturation compensator.
[0021] Furthermore, the dynamic model of the flexible solar panel of the flexible fault satellite is as follows:
[0022]
[0023] in, The unit mass matrix, For the system's generalized coordinates, Generalized external force matrix Damping force matrix This is the elastic force matrix.
[0024] Furthermore, the process of establishing the dynamic model of the flexible solar panel of the flexible fault satellite is as follows:
[0025] The flexible solar panel of the flexible fault satellite is described by a rectangular four-node element. The kinetic energy of the flexible solar panel element is calculated by using the displacement field function of any point in the flexible solar panel element in the global coordinate system.
[0026] Based on Kirchhoff's plate theory, the strain energy of the flexible solar panel is calculated, which includes the longitudinal shear strain of the mid-surface and the bending and torsion of the mid-surface.
[0027] Based on the kinetic and strain energies of the flexible solar panel unit, a dynamic model of the flexible solar panel is derived using the Euler-Lagrange equations.
[0028] Furthermore, the dynamic equations of the flexible fault satellite rigid-flexible coupling system are as follows:
[0029]
[0030] in, Indicates system quality. Represents the system's generalized coordinates. Represents the Lagrange multiplier vector. Represent the constraint equations. It represents the generalized force vector.
[0031] The predefined time-contact despinning control system for flexible fault satellites includes:
[0032] A modeling module is used to acquire the position and attitude of the servicing spacecraft and the flexible fault satellite, and to establish a dynamic model of the servicing spacecraft and the flexible fault satellite.
[0033] The design module is used to design a predefined time based on adaptive compensation, according to the state of the servicing spacecraft and the interference estimate.
[0034] The control module is used to establish a dynamic model of the flexible solar panel of the flexible fault satellite. Based on the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish a dynamic model of the rigid-flexible coupled system of the flexible fault satellite. Based on the state of the rigid-flexible coupled system of the flexible fault satellite, contact despin control is performed.
[0035] A terminal device includes a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the steps of the method.
[0036] A computer-readable storage medium storing a computer program that, when executed by a processor, implements the steps of the method.
[0037] Compared with the prior art, the present invention has the following beneficial effects:
[0038] This invention provides a predefined-time contact despin control method for a flexible faulty satellite. By acquiring the positions and attitudes of both the servicing spacecraft and the flexible faulty satellite, dynamic models of both are established. Based on the state of the servicing spacecraft and disturbance estimates, a predefined time based on adaptive compensation is designed. By constructing a novel disturbance observer and adaptive compensation framework, the complexity of the control algorithm is reduced while significantly improving the system's anti-interference capability and ensuring precise control performance. This provides a solution with high real-time performance and strong robustness for tasks such as despinning of non-cooperative space targets, achieving robustness of the servicing spacecraft under strong disturbances. Furthermore, a dynamic model of the flexible solar panel of the flexible faulty satellite is established. Based on the predefined time and the state of the flexible faulty satellite, a dynamic model of the rigid-flexible coupled system of the flexible faulty satellite is established using the Lagrange multiplier method. Contact despin control is then performed based on the state of the rigid-flexible coupled system of the flexible faulty satellite, ensuring the servicing spacecraft achieves predefined-time convergence and ensuring efficient and effective despinning of the flexible faulty satellite. This invention features a simple structure, reducing computational complexity. Fixed-time convergence can be achieved by adjusting only one parameter, avoiding tedious parameter tuning. Therefore, it can provide spacecraft stability within a fixed time frame based on actual despinning tasks, easily meeting the requirements of real-time despinning missions. The invention's predefined-time robust control method introduces a time-varying gain mechanism, overcoming the theoretical limitations of traditional fixed-time control. It achieves engineering-oriented, precise control of the preset time, allowing direct setting of the system's convergence time upper bound according to mission requirements. This enables preset-time stable control of spacecraft and high-precision despinning operation of flexible fault satellites under conditions of strong time-varying interference and uncertainty, significantly improving the engineering applicability of on-orbit operations and demonstrating strong engineering application value. Attached Figure Description
[0039] To more clearly illustrate the technical solutions of the embodiments of the present invention, the accompanying drawings used in the embodiments will be briefly introduced below. It should be understood that the following drawings only show some embodiments of the present invention and should not be regarded as a limitation on the scope. For those skilled in the art, other related drawings can be obtained based on these drawings without creative effort.
[0040] Figure 1 This is a flowchart of the predefined time contact despin control method for flexible fault satellites according to the present invention.
[0041] Figure 2This is a schematic diagram of the contact despinning process of the flexible fault satellite of the present invention.
[0042] Figure 3 This is a diagram illustrating the attitude change process of a spacecraft used in this invention.
[0043] Figure 4 This is a diagram illustrating the thrust variation process of a spacecraft used in this invention.
[0044] Figure 5 A graph showing the change in angular velocity of the spacecraft used in this invention.
[0045] Figure 6 This is a graph showing the change history of the angular velocity of the flexible fault satellite according to the present invention.
[0046] Figure 7 This is a graph showing the change in the position of the flexible fault satellite according to the present invention.
[0047] Figure 8 This is a translational velocity history diagram of the flexible fault satellite of the present invention.
[0048] Figure 9 This is a schematic diagram of the predefined time contact despin control system for a flexible fault satellite, which is a preferred embodiment of the present invention.
[0049] Figure 10 This is a schematic diagram of the electronic device structure according to a preferred embodiment of the present invention. Detailed Implementation
[0050] The following description, in conjunction with the accompanying drawings, illustrates exemplary embodiments of this application, including various details to aid understanding. These should be considered merely exemplary. Therefore, those skilled in the art will recognize that various changes and modifications can be made to the embodiments described herein without departing from the scope and spirit of this application. Similarly, for clarity and brevity, descriptions of well-known functions and structures are omitted in the following description.
[0051] Obviously, the described embodiments are only some, not all, of the embodiments in this application. All other embodiments obtained by those skilled in the art based on the embodiments in this application without inventive effort are within the scope of protection of this application.
[0052] It should be noted that the terminals involved in the embodiments of this 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 (e.g., smart glasses, smartwatches, smart bracelets, etc.), smart home devices, and other smart devices.
[0053] Furthermore, the term "and / or" in this article is merely a description of the relationship between related objects, indicating that three relationships can exist. For example, A and / or B can represent: A existing alone, A and B existing simultaneously, or B existing alone. Additionally, the character " / " in this article generally indicates that the preceding and following related objects have an "or" relationship.
[0054] The present invention will now be described in further detail with reference to the accompanying drawings:
[0055] See Figure 1 This invention provides a predefined time-contact despin control method for flexible fault satellites, specifically including the following steps:
[0056] Step 1: Obtain the position and attitude of the servicing spacecraft and establish a dynamic model of the servicing spacecraft.
[0057] The service spacecraft position and attitude dynamics equations are expressed as follows:
[0058] (1)
[0059] (2)
[0060] In the formula: Indicates the mass of the spacecraft base. This represents the position vector from the center of mass of the serving spacecraft to the origin of the inertial frame. Indicates the moment of inertia of the service spacecraft. Represents the Euler angle vector of the servicing spacecraft. This represents the attitude angular velocity vector of the servicing spacecraft. This represents the control torque acting on the service spacecraft matrix. Control forces acting on the service spacecraft matrix, matrix Represented as
[0061] (3)
[0062] For ease of establishing the despin control system, the dynamic equations are written as an Euler-Langargen system, expressed as follows:
[0063] (4)
[0064] In the formula: , .
[0065] Step 2: Based on the state of the servicing spacecraft and the disturbance estimate, design a predefined time control based on adaptive compensation.
[0066] Redefine the service spacecraft status as and The dynamics system of a spacecraft can be written in the following state-space form.
[0067] (5)
[0068] In the formula: , , This represents the coupled disturbance acting on the spacecraft system, where and Let be the Jacobian matrices, representing the constraint equations with respect to attitude variables. and position variables The partial derivatives of .
[0069] Tracking error is defined as
[0070] (6)
[0071] In the formula, and Let be the desired trajectory and its first derivative.
[0072] Step 1: Define the backstep variable as follows
[0073] (7)
[0074] In the formula: Designed as a virtual control input
[0075] (8)
[0076] In the formula, , , , This indicates the preset time.
[0077] The first derivative with respect to time is obtained
[0078] (9)
[0079] In the formula: , ,
[0080] , This refers to the state of the anti-saturation compensator.
[0081] Define the first Lyapunov function
[0082] (10)
[0083] Taking the first derivative of equation (9) with respect to time, we can obtain...
[0084] (11)
[0085] Substituting the virtual control input into equation (10) yields
[0086] (12)
[0087] In the formula: It is a positive number.
[0088] Step 2: Define the second Lyapunov function
[0089] (13)
[0090] Taking the first derivative of the second equation (7) with respect to time, we can obtain...
[0091] (14)
[0092] Taking the derivative of equation (13) with respect to time, we can obtain...
[0093] (15)
[0094] The actual control input can be designed as
[0095] (16)
[0096] (17)
[0097] In the formula: For the positive constants to be designed.
[0098] Interference estimation algorithm design
[0099] (18)
[0100] In the formula: This is the observer gain.
[0101] The estimation algorithm is designed as follows
[0102] (19)
[0103] In the formula: express Estimation error.
[0104] The interference estimate is designed as follows
[0105] (20)
[0106] The error dynamics are
[0107] (twenty one)
[0108] Substituting equation (15) into equation (16) yields
[0109] (twenty two)
[0110] In the formula: and It is a positive number.
[0111] Step 3: System stability analysis.
[0112] Consider the following Lyapunov function
[0113] (twenty three)
[0114] Taking the first derivative of equation (23) with respect to time, we can obtain...
[0115] (twenty four)
[0116] In the formula: .
[0117] We can conclude that: , and At the preset time The convergence occurs within a smaller region of the equilibrium point. Therefore, the service spacecraft state tracks the desired state over a preset time, and the designed disturbance observer estimates the coupled disturbances over the preset time.
[0118] Step 4: Establish dynamic models of the flexible fault satellite and the flexible solar panel. Based on the predefined time and the state of the flexible fault satellite, use the Lagrange multiplier method to establish a dynamic model of the rigid-flexible coupled system of the flexible fault satellite. Based on the state of the rigid-flexible coupled system of the flexible fault satellite, perform contact despin control.
[0119] A flexible solar panel dynamic model based on the absolute nodal coordinate method was established using rectangular four-node elements to describe the flexible plate. Each element contains four nodes, and the coordinate vector of each node has nine components, resulting in a total of 36 nodal coordinate vectors and 36 degrees of freedom for the element. Any point within the flexible solar panel element... The displacement field function in the global coordinate system can be expressed as:
[0120] (25)
[0121] In the formula: Let be a shape function, where, , , , , , , , , , , , , , For unit length, For unit width, These are the node coordinates of the element.
[0122] Let the coordinate array of any node in the flexible solar panel element be represented as:
[0123] (26)
[0124] In the formula: Let A be the global position vector. for For local coordinates The bias leads to for For local coordinates The partial derivatives, Let B be the global position vector. for For local coordinates The bias leads to for For local coordinates The partial derivatives, Let C be the global position vector. for For local coordinates The bias leads to for For local coordinates The partial derivatives, Let D be the global position vector. for For local coordinates The bias leads to for For local coordinates The partial derivatives of .
[0125] but Represented as
[0126] (27)
[0127] In the formula: , , , These are the node coordinates of unit nodes A, B, C, and D, respectively.
[0128] Unit kinetic energy is expressed as
[0129] (28)
[0130] In the formula: Represents the unit mass matrix, For unit density, The unit thickness is denoted as .
[0131] According to Kirchhoff's plate theory, the strain energy of a sail is generated by two parts: one part is generated by the longitudinal shear strain of the mid-surface, and the other part is generated by the bending and torsion of the mid-surface.
[0132] The strain vector of the mid-surface is expressed as
[0133] (29)
[0134] In the formula: , , , , .
[0135] The curvature vector of the mid-surface is represented as
[0136] (30)
[0137] In the formula: , , , , , , , .
[0138] The strain energy of the solar panel element is expressed as:
[0139] (31)
[0140] In the formula: , , , , , , , , , , , , , , . and The elasticity coefficient matrix is represented as follows:
[0141] (32)
[0142] In the formula: For elastic modulus, It is Poisson's ratio.
[0143] Elastic force calculation is expressed as
[0144] (33)
[0145] Considering the viscoelasticity of the material, the element shear strain and the generalized forces generated by bending and torsion are expressed as follows:
[0146] (34)
[0147] In the formula: and This represents the corresponding damping coefficient.
[0148] Using the Euler-Lagrange equations, the dynamic equations of the solar panel system can be derived as follows:
[0149] (35)
[0150] in, The unit mass matrix, For the system's generalized coordinates, Generalized external force matrix Damping force matrix This is the elastic force matrix.
[0151] The satellite attitude dynamics equation is expressed as follows:
[0152] (36)
[0153] In the formula: Indicates quality, For position vectors, Represents the Euler angle vector. Represents the angular velocity vector. express antisymmetric matrix, Represented as
[0154] (37)
[0155] The dynamic equations of the flexible satellite rigid-flexible coupling system are established using the Lagrange multiplier method. These dynamic equations are expressed as a system of differential-algebraic equations of index-3:
[0156] (38)
[0157] In the formula: For system quality, For the system's generalized coordinates, For Lagrange multiplier vectors, For the constraint equations, It represents the generalized force vector.
[0158] To facilitate solving the above equations, they are usually transformed into ordinary differential equations with index -1, and the Baumgarte stability method is used to solve the system of equations:
[0159] (39)
[0160] In the formula: , , , and Feedback control parameters are used to limit violations of speed and position constraints.
[0161] The present invention will be further described in detail below through specific embodiments:
[0162] Example 1:
[0163] In this embodiment, the predefined time-contact despin control method for flexible fault satellites specifically includes the following steps:
[0164] Step 1: Obtain the position and attitude of the servicing spacecraft and establish a dynamic model of the servicing spacecraft.
[0165] The service spacecraft position and attitude dynamics equations are expressed as follows:
[0166] (1)
[0167] In the formula: Indicates the mass of the base. This represents the position vector from the center of mass of the serving spacecraft to the origin of the inertial frame. Indicates the moment of inertia of a spacecraft. Represents the Euler angle vector, used to describe the attitude of a spacecraft. This represents the spacecraft's attitude angular velocity vector. The matrix represents the control torque acting on the matrix. Represented as
[0168] (2)
[0169] The attitude dynamics equations can be written in the following Euler-Langrangian form, and the service spacecraft attitude dynamics model can be represented as follows:
[0170] (3)
[0171] In the formula: , , This represents the position vector from the center of mass of the serving spacecraft to the origin of the inertial frame. Control forces acting on the matrix.
[0172] Step 2: Based on the state of the servicing spacecraft and the disturbance estimate, design a predefined time control based on adaptive compensation.
[0173] Redefine the service spacecraft status as and The dynamics system of a spacecraft can be written in the following state-space form.
[0174] (4)
[0175] In the formula: , , This represents the coupled disturbance acting on the spacecraft system, where and Let be the Jacobian matrices, representing the constraint equations with respect to attitude variables. and position variables The partial derivatives of .
[0176] (5)
[0177] Among them, virtual control input Designed for
[0178] (6)
[0179] In the formula: , , , This indicates the preset time.
[0180] The actual control input can be designed as
[0181] (7)
[0182] (8)
[0183] In the formula: For the positive constants to be designed.
[0184] Interference estimation algorithm design
[0185] (9)
[0186] In the formula: This is the observer gain.
[0187] The estimation algorithm is designed as follows
[0188] (10)
[0189] In the formula: express Estimation error.
[0190] The interference estimate is designed as follows
[0191] (11)
[0192] Step 3: Establish a dynamic model of the flexible fault satellite and the flexible solar panel. Based on the predefined time and the state of the flexible fault satellite, use the Lagrange multiplier method to establish a dynamic model of the rigid-flexible coupled system of the flexible fault satellite. Based on the state of the rigid-flexible coupled system of the flexible fault satellite, perform contact despin control, such as... Figure 2 As shown.
[0193] A flexible solar panel was described using rectangular four-node elements, and a dynamic model of the flexible solar panel based on the absolute nodal coordinate method was established. Each element contains four nodes, and the coordinate vector of each node has nine components, resulting in a total of 36 nodal coordinate vectors and 36 degrees of freedom for the element. Any point within the flexible solar panel element... The displacement field function in the global coordinate system can be expressed as:
[0194] (12)
[0195] In the formula: Let be a shape function, where,
[0196] , , , , , , , , , , , , , For unit length, For unit width, These are the node coordinates of the element.
[0197] Let the coordinate array of any node in the windsurfing element be... Represented as
[0198] (13)
[0199] In the formula: , , , These are the node coordinates of unit nodes A, B, C, and D, respectively.
[0200] Unit kinetic energy is expressed as
[0201] (14)
[0202] In the formula: Represents the unit mass matrix, For unit density, The unit thickness is denoted as .
[0203] According to Kirchhoff's plate theory, the strain energy of a sail is generated by two parts: one part is generated by the longitudinal shear strain of the mid-surface, and the other part is generated by the bending and torsion of the mid-surface.
[0204] The curvature vector of the mid-surface is represented as
[0205] (15)
[0206] In the formula: , , , , , , , .
[0207] The strain energy of the solar panel element is expressed as:
[0208] (16)
[0209] In the formula: , , , , , , , , , , , , , , . and The elasticity coefficient matrix is represented as follows:
[0210] (17)
[0211] In the formula: It is the elastic modulus; It is Poisson's ratio.
[0212] Elastic force calculation is expressed as
[0213] (18)
[0214] Considering the viscoelasticity of the material, the element shear strain and the generalized forces generated by bending and torsion are expressed as follows:
[0215] (19)
[0216] In the formula: and This represents the corresponding damping coefficient.
[0217] Using the Euler-Lagrange equations, the dynamic equations of the solar panel system can be derived as follows:
[0218] (20)
[0219] in, The unit mass matrix, For the system's generalized coordinates, Generalized external force matrix Damping force matrix This is the elastic force matrix.
[0220] The satellite attitude dynamics equation is expressed as follows:
[0221] (twenty one)
[0222] In the formula: Indicates quality, For position vectors, Represents the Euler angle vector. Represents the angular velocity vector. express antisymmetric matrix, Represented as
[0223] (twenty two)
[0224] The dynamic equations of the flexible satellite rigid-flexible coupling system are established using the Lagrange multiplier method. These dynamic equations are expressed as a system of differential-algebraic equations of index-3:
[0225] (twenty three)
[0226] In the formula: For system quality, For the system's generalized coordinates, For Lagrange multiplier vectors, For the constraint equations, It represents the generalized force vector.
[0227] To facilitate solving the above equations, they are usually transformed into ordinary differential equations with index -1, and the Baumgarte stability method is used to solve the system of equations:
[0228] (twenty four)
[0229] In the formula: , , , and Feedback control parameters are used to limit violations of speed and position constraints.
[0230] Step 4: Perform numerical simulation.
[0231] The parameters selected for the servicing spacecraft dynamics model are: moment of inertia. ,quality The initial value is set as follows: ; ; ; .
[0232] The dynamic simulation parameters of the flexible fault satellite are shown in Tables 1 and 2:
[0233] Table 1 Physical parameters of flexible satellite solar panels
[0234]
[0235] Table 2 Physical parameters of the rigid body at the center of the flexible satellite
[0236]
[0237] The controller parameters are selected as follows: ; ; ; ; .
[0238] Numerical simulation results are as follows Figures 3-8 As shown. Figure 3 The graph showing the change in the attitude angle of the servicing spacecraft is presented. It can be seen that the servicing spacecraft went through 9 despinning processes, and the attitude angle converged to the equilibrium position after each despinning process. Figure 4 The evolution of the thruster is shown, and it can be seen that the thrust is limited to the set saturation range during each despinning process. Figure 5 The graph depicts the evolution of the angular velocity of the service spacecraft, showing that the angular velocity can quickly converge to the equilibrium position to maintain the desired orientation of the control sticks.
[0239] Figure 6 The evolution of the satellite's angular velocity is shown, revealing that after nine despinning processes, the satellite's angular velocity decreases to a range that satisfies subsequent acquisition requirements. Figure 7 The change in the satellite's position was plotted, showing that the satellite's position in the X and X directions gradually moved away from each other, which is due to the effect of contact forces. Figure 8 The evolution of the satellite's translational velocity is presented, showing that the translational velocity exhibits an increasing trend and oscillations.
[0240] This invention also provides a predefined time-contact despinning control system for flexible fault satellites, such as... Figure 9 As shown, the system includes: a modeling module, a design module, and a control module.
[0241] A modeling module is used to acquire the position and attitude of the servicing spacecraft and the flexible fault satellite, and to establish a dynamic model of the servicing spacecraft and the flexible fault satellite.
[0242] The design module is used to design a predefined time based on adaptive compensation, according to the state of the servicing spacecraft and the interference estimate.
[0243] The control module is used to establish a dynamic model of the flexible solar panel of the flexible fault satellite. Based on the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish a dynamic model of the rigid-flexible coupled system of the flexible fault satellite. Based on the state of the rigid-flexible coupled system of the flexible fault satellite, contact despin control is performed.
[0244] It is understood that the predefined time contact despinning control system for flexible fault satellites provided by the present invention corresponds to the predefined time contact despinning control method for flexible fault satellites provided in the foregoing embodiments. The relevant technical features of the predefined time contact despinning control system for flexible fault satellites can be referred to the relevant technical features of the predefined time contact despinning control method for flexible fault satellites, and will not be repeated here.
[0245] Another object of the present invention is to provide an electronic device, such as... Figure 10 As shown, it includes a memory, a processor, and a computer program stored in the memory and executable on the processor, the processor performing the steps of the predefined time-contact despin control method for the flexible fault satellite.
[0246] The predefined time-contact despin control method for the flexible fault satellite includes the following steps:
[0247] Obtain the position and attitude of the servicing spacecraft and the flexible fault satellite, and establish dynamic models of the servicing spacecraft and the flexible fault satellite;
[0248] Based on the state of the servicing spacecraft and the interference estimate, a predefined time based on adaptive compensation is designed.
[0249] A dynamic model of the flexible solar panel of the flexible fault satellite is established. Based on the predefined time and the state of the flexible fault satellite, a dynamic model of the rigid-flexible coupled system of the flexible fault satellite is established using the Lagrange multiplier method. Contact despin control is then performed based on the state of the rigid-flexible coupled system of the flexible fault satellite.
[0250] A fourth objective of this invention is to provide a computer-readable storage medium storing a computer program that, when executed by a processor, implements the steps of a predefined time-contact despin control method for a flexible faulty satellite.
[0251] The predefined time-contact despin control method for the flexible fault satellite includes the following steps:
[0252] Obtain the position and attitude of the servicing spacecraft and the flexible fault satellite, and establish dynamic models of the servicing spacecraft and the flexible fault satellite;
[0253] Based on the state of the servicing spacecraft and the interference estimate, a predefined time based on adaptive compensation is designed.
[0254] A dynamic model of the flexible solar panel of the flexible fault satellite is established. Based on the predefined time and the state of the flexible fault satellite, a dynamic model of the rigid-flexible coupled system of the flexible fault satellite is established using the Lagrange multiplier method. Contact despin control is then performed based on the state of the rigid-flexible coupled system of the flexible fault satellite.
[0255] Those skilled in the art will understand that embodiments of the present invention can be provided as methods, systems, or computer program products. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention can take the form of a computer program product embodied 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.
[0256] This invention is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.
[0257] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.
[0258] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.
[0259] 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 it. Although the present invention has been described in detail with reference to the above embodiments, those skilled in the art should understand that modifications or equivalent substitutions can still be made to the specific implementation of the present invention. Any modifications or equivalent substitutions that do 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 despin control method for flexible fault satellites, characterized in that, include: Obtain the position and attitude of the servicing spacecraft and the flexible fault satellite, and establish dynamic models of the servicing spacecraft and the flexible fault satellite; Based on the state of the servicing spacecraft and the interference estimate, a predefined time based on adaptive compensation is designed. A dynamic model of the flexible solar panel of the flexible fault satellite is established. Based on the predefined time and the state of the flexible fault satellite, a dynamic model of the rigid-flexible coupled system of the flexible fault satellite is established using the Lagrange multiplier method. Contact despin control is then performed based on the state of the rigid-flexible coupled system of the flexible fault satellite. The process of establishing the dynamic model of the flexible solar panel of the flexible fault satellite is as follows: The flexible solar panel of the flexible fault satellite is described by a rectangular four-node element. The kinetic energy of the flexible solar panel element is calculated by using the displacement field function of any point in the flexible solar panel element in the global coordinate system. Based on Kirchhoff's plate theory, the strain energy of the flexible solar panel is calculated, which includes the longitudinal shear strain of the mid-surface and the bending and torsion of the mid-surface. Based on the kinetic energy and strain energy of the flexible solar panel unit, the dynamic model of the flexible solar panel is derived using the Euler-Lagrange equations. The dynamic equations of the flexible fault satellite rigid-flexible coupling system are as follows: in, Indicates system quality. Represents the system's generalized coordinates. Represents the Lagrange multiplier vector. Represent the constraint equations. It represents the generalized force vector.
2. The predefined time contact despin control method for flexible faulty satellites according to claim 1, characterized in that, The dynamic model of the service spacecraft is as follows: in, Indicates the mass of the spacecraft base. This represents the position vector from the center of mass of the serving spacecraft to the origin of the inertial frame. Indicates the moment of inertia of the service spacecraft. Represents the Euler angle vector of the servicing spacecraft. This represents the attitude angular velocity vector of the servicing spacecraft. This represents the control torque acting on the service spacecraft matrix. Control forces acting on the service spacecraft matrix, matrix Represented as .
3. The predefined time contact despin control method for flexible faulty satellites according to claim 1, characterized in that, The dynamic model of the flexible fault satellite is as follows: in, Indicates the mass of a satellite with a flexible fault. For the position vector of the flexible fault satellite, The Euler angle vector representing a satellite with a flexible fault. The vector representing the angular velocity of a satellite with a flexible fault. express antisymmetric matrix, Represented as .
4. The predefined time contact despin control method for flexible faulty satellites according to claim 1, characterized in that, The interference estimate is: in, For observer gain, This indicates the state of the anti-saturation compensator.
5. The predefined time contact despin control method for flexible faulty satellites according to claim 1, characterized in that, The dynamic model of the flexible solar panel of the aforementioned flexible fault satellite is as follows: in, The unit mass matrix, For the system's generalized coordinates, Generalized external force matrix Damping force matrix This is the elastic force matrix.
6. A predefined time-contact despin control system for a flexible fault satellite, characterized in that, The steps for implementing the method of claim 1 include: A modeling module is used to acquire the position and attitude of the servicing spacecraft and the flexible fault satellite, and to establish a dynamic model of the servicing spacecraft and the flexible fault satellite. The design module is used to design a predefined time based on adaptive compensation, according to the state of the servicing spacecraft and the interference estimate. The control module is used to establish a dynamic model of the flexible solar panel of the flexible fault satellite. Based on the predefined time and the state of the flexible fault satellite, the Lagrange multiplier method is used to establish a dynamic model of the rigid-flexible coupled system of the flexible fault satellite. Based on the state of the rigid-flexible coupled system of the flexible fault satellite, contact despin control is performed.
7. 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, it implements the steps of the method according to any one of claims 1 to 5.
8. A computer-readable storage medium storing a computer program, characterized in that, When the computer program is executed by a processor, it implements the steps of the method according to any one of claims 1 to 5.