A method and system for cooperative electromagnetic takeover of a large-fault spacecraft

Abstract

The application discloses a kind of methods and systems for large-scale fault spacecraft's cooperative electromagnetic takeover, it is related to spacecraft electromagnetic takeover technical field, wherein the method comprises: determining reference position trajectory based on task demand;Release multiple service spacecraft to reach initial takeover position;Based on the relative information of service spacecraft and fault spacecraft, obtain the corresponding disturbance estimate value;Keep the original motion direction unchanged, adjust the relative position direction of service spacecraft, so that it is consistent with the relative position direction of service spacecraft that can receive instruction, obtain rotating relative position;Based on rotating relative position, reference position trajectory and disturbance estimate value, obtain control input;Based on control input, each service spacecraft and fault spacecraft realize docking;Based on the service spacecraft that can receive instruction receives control information and distributes to the actuator of all service spacecraft, realize the control of fault spacecraft. Reduce on-board computing resources, realize accurate compliant takeover control with fault spacecraft.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of spacecraft electromagnetic docking, in particular to a cooperative electromagnetic docking method and system for large-scale fault spacecraft. BACKGROUND

[0002] According to incomplete statistics, more than 17000 spacecraft have been launched and entered the designated orbit worldwide, and 35% of them have become fault spacecraft due to mechanical collision, energy depletion, part failure and other reasons, which are potential "time bombs" in space and threaten the safety of normally operating spacecraft. At present, the fault of the attitude and orbit control related system is the most common fault of spacecraft. The effective payload of the fault spacecraft is usually intact and can continue to work, and if the attitude and orbit control ability of the fault spacecraft can be restored through docking, it is a feasible solution. The service spacecraft and the fault spacecraft are connected by electromagnetic docking, and the electromagnetic force is used as the control force to realize soft docking, so as to realize the attitude and orbit control of the fault spacecraft

[0003] However, in the process of electromagnetic docking, the origin of the coordinate system is usually selected at the center of mass of the docking system, not at the center of mass of the fault spacecraft. Since the center of mass of the system will change slowly with the docking process, deviation will be caused, and then the docking accuracy will be affected. In addition, for each service spacecraft docking with the fault spacecraft, the trajectory data must be pre-stored for tracking control during design, which occupies part of the on-board computing resources.

[0004] Therefore, how to realize accurate soft docking control with the fault spacecraft while keeping the center of mass stable under the premise of reducing on-board computing resources is a problem to be solved by those skilled in the art. SUMMARY

[0005] Therefore, the present application provides a cooperative electromagnetic docking method and system for large-scale fault spacecraft, which realizes accurate soft docking control with the fault spacecraft while keeping the center of mass stable under the premise of reducing on-board computing resources.

[0006] In order to achieve the above purpose, the present application adopts the following technical solutions:

[0007] A cooperative electromagnetic docking method for large-scale fault spacecraft, comprising:

[0008] determining a reference position trajectory based on task requirements;

[0009] releasing a plurality of service spacecraft to arrive at an initial docking position;

[0010] obtaining a corresponding disturbance estimation value based on the relative information of the service spacecraft and the fault spacecraft;

[0011] keeping the original moving direction unchanged, adjusting the relative position direction of the service spacecraft to be consistent with the relative position direction of the service spacecraft capable of receiving the instruction, to obtain a rotating relative position;

[0012] based on the rotating relative position, the reference position trajectory and the disturbance estimation value, obtaining a control input;

[0013] based on the control input, each of the service spacecraft and the failure spacecraft realizes docking;

[0014] based on the service spacecraft capable of receiving the instruction, receiving control information and distributing to the actuators of all service spacecraft, realizing the control of the failure spacecraft.

[0015] Preferably, the initial takeover position determination method is:

[0016] based on the center of mass of the system composed of multiple service spacecraft and the failure spacecraft coinciding with the center of mass of the failure spacecraft, at this time, multiple service spacecraft arrive at the initial takeover position.

[0017] Preferably, the disturbance estimation value acquisition method is:

[0018] obtaining the relative position of the service spacecraft and the failure spacecraft;

[0019] based on the relative position, obtaining the relative velocity of the service spacecraft and the failure spacecraft by derivation;

[0020] based on the relative position and the relative velocity, inputting into the disturbance observer to obtain the disturbance estimation value;

[0021] the relative position and the relative velocity are the relative information.

[0022] Preferably, the control input acquisition method is:

[0023] based on the rotating relative position and the reference position trajectory, obtaining a relative position error;

[0024] based on the relative position error and the communication topology relationship between the service spacecraft, obtaining a relative information error with other service spacecraft;

[0025] based on the relative information error, obtaining a control force;

[0026] based on the control force and the corresponding disturbance estimation value, obtaining the control input.

[0027] Preferably, the relative position error acquisition method is:

[0028] a mass coefficient based on a ratio of a mass of the service spacecraft to a mass of the service spacecraft capable of receiving the command;

[0029] a transformed relative position based on the rotational relative position and the corresponding mass coefficient;

[0030] the relative position error based on the transformed relative position and the reference position trajectory.

[0031] Preferably, the method for obtaining the control force comprises:

[0032] a sliding mode surface based on a derivative of the rotational relative position and the relative information error;

[0033] the adjusted control force based on the sliding mode surface, the derivative of the rotational relative position, the rotational relative position, the derivative of the relative information error and the relative information error.

[0034] Preferably, the control of the faulty spacecraft is implemented, and specifically comprises:

[0035] based on the communication conditions between each of the service spacecraft and the usage conditions of the actuators of each of the service spacecraft, using graph theory knowledge, taking each of the actuators as a node, and abstracting the communication conditions between each of the service spacecraft as edges connecting the nodes;

[0036] based on the edges, assigning corresponding influence coefficients;

[0037] based on the output torque, the thruster life and the energy consumption, setting a comprehensive factor for each of the nodes;

[0038] based on the comprehensive factor, the control information and the influence coefficients, constructing a torque distribution objective function containing constraints;

[0039] based on the torque distribution objective function, solving to obtain the control force and the control torque on each of the actuators;

[0040] based on the control force and the control torque, implementing the control of the faulty spacecraft.

[0041] Preferably, the comprehensive factor specifically comprises:

[0042] based on the accuracy of the output torque and the saturation phenomenon of the flywheel actuator, setting an ability factor;

[0043] based on the thruster life, setting a life factor;

[0044] based on the energy consumption, setting a performance index;

[0045] The comprehensive factor is composed of the capability factor, the life factor and the performance index.

[0046] Preferably, the moment distribution target function is specifically:

[0047]

[0048] s.t.Ak i =τ d ;

[0049] A=[D ij ]∈R n×n ;

[0050] Wherein, β 1i , β 2i and β 3i represent weight coefficients, g ωi represents a capability factor, g ni represents a life factor, g εi represents a performance index, k i represents an output of the ith actuator, τ d represents control information, D ij represents an influence coefficient, and n represents the number of service spacecrafts.

[0051] A cooperative electromagnetic takeover system for large-scale fault spacecrafts, comprising: a trajectory determination module, an initial position determination module, a disturbance estimation module, a direction adjustment module, a final docking module and a control module;

[0052] The trajectory determination module is configured to determine a reference position trajectory based on task requirements.

[0053] The initial position determination module is configured to release a plurality of service spacecrafts to arrive at an initial takeover position.

[0054] The disturbance estimation module is configured to obtain a corresponding disturbance estimation value based on relative information of the service spacecrafts and the fault spacecraft.

[0055] The direction adjustment module is configured to keep the original movement direction unchanged, adjust a relative position direction of the service spacecrafts to be consistent with a relative position direction of the service spacecrafts capable of receiving instructions, and obtain a rotated relative position.

[0056] The final docking module is configured to obtain a control input based on the rotated relative position, the reference position trajectory and the disturbance estimation value, and implement docking of each of the service spacecrafts and the fault spacecraft based on the control input.

[0057] The control module is used for receiving control information by the service spacecraft capable of receiving instructions and distributing to all execution mechanisms of the service spacecraft, so as to realize takeover control of the fault spacecraft.

[0058] Compared with the prior art, the application provides a cooperative electromagnetic takeover method and system for large-scale fault spacecraft, which uses electromagnetic force as control force to effectively reduce the impact force of contact type takeover and avoid secondary damage, realizes takeover by using electromagnetic force, and has the advantages of high fault tolerance, no rigid collision, no debris generation, no plume pollution, flexibility, wide application range and the like; the application only has one service spacecraft capable of receiving instructions, the rest of the service spacecraft perform electromagnetic docking according to information between each other and keep the center of mass of the system composed of each service spacecraft and the fault spacecraft unchanged, improves the accuracy of docking, and reduces the memory requirement; after the docking is completed, the application uses the execution mechanisms of each service spacecraft according to the designed torque distribution rule, so as to realize accurate control of the attitude and orbit of the fault spacecraft.

[0059] The motor driving control algorithm is applied to the takeover control of large-scale spacecraft, and the combined spacecraft after takeover is controlled in attitude and orbit. First, the multiple service spacecraft are connected with the fault spacecraft in a cooperative electromagnetic docking manner to form a combined spacecraft. There are a large number of nonlinear parts in the attitude and orbit model of the combined spacecraft, the motor driving control algorithm is used to study the nonlinear relationship of the controlled object, a nonlinear model of the attitude and orbit coupling dynamics equation is established, a feedback compensation method is used to improve the response speed and stability of the motor driving execution mechanism for the control instruction delay caused by nonlinearity, the control precision of the combined spacecraft is improved. BRIEF DESCRIPTION OF DRAWINGS

[0060] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the drawings needed to be used in the embodiments or prior art description will be briefly introduced as follows. Obviously, the drawings in the following description only constitute the embodiments of the present application, and for those skilled in the art, other drawings can be obtained without creative labor on the basis of the provided drawings.

[0061] Figure 1 A cooperative electromagnetic takeover method for large-scale fault spacecraft is provided.

[0062] Figure 2 A control input acquisition method is provided.

[0063] Figure 3 A cooperative electromagnetic takeover system structure schematic diagram for large-scale fault spacecraft is provided.

[0064] Figure 4 The structural block diagram of the computer device provided by the present application is shown. DETAILED DESCRIPTION

[0065] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work fall within the protection scope of the present application.

[0066] Embodiment 1

[0067] As shown in the accompanying drawings, Figure 1 the embodiment of the present application discloses a cooperative electromagnetic takeover method for large-fault spacecraft, comprising:

[0068] determining a reference position trajectory based on task requirements;

[0069] releasing a plurality of service spacecraft to arrive at an initial takeover position;

[0070] obtaining a corresponding disturbance estimation value based on the relative information of the service spacecraft and the fault spacecraft;

[0071] keeping the original motion direction unchanged, adjusting the relative position direction of the service spacecraft to be consistent with the relative position direction of the service spacecraft capable of receiving instructions, and obtaining a rotating relative position;

[0072] obtaining a control input based on the rotating relative position, the reference position trajectory and the disturbance estimation value;

[0073] based on the control input, each service spacecraft and the fault spacecraft realize docking;

[0074] based on the service spacecraft capable of receiving instructions receiving control information and distributing to the actuators of all service spacecraft, the control of the fault spacecraft is realized.

[0075] Embodiment 2

[0076] There are too many fault spacecraft in the space orbit today, which seriously endangers the safety of other spacecraft. In order to restore the attitude and orbit control of the fault spacecraft, a plurality of service spacecraft with different functions and sizes can be fixed with the fault spacecraft, and communication, observation or different sizes of control force and control moment can be generated according to functions.

[0077] The electromagnetic docking mode is used to connect the service spacecraft with the fault spacecraft, and the electromagnetic force can be used as the control force to realize the compliant docking in space, and the electromagnetic force is a non-contact field force, which has the advantages of continuity, reversibility and synchronous controllability, and compared with other contact pipe connection methods, the impact force is effectively reduced, and greater damage to the spacecraft is avoided.

[0078] Based on this, the embodiment of the application discloses a cooperative electromagnetic pipe connection method for large fault spacecraft, comprising:

[0079] The reference position trajectory is determined based on the task demand.

[0080] Preferably, in the embodiment, the reference position trajectory η d is the trajectory of the electromagnetic docking displacement with respect to time.

[0081] The multiple service spacecraft are released to reach the initial pipe connection position.

[0082] Preferably, the service spacecraft include a service spacecraft capable of receiving instructions and multiple service spacecraft, wherein the service spacecraft capable of receiving instructions can receive relevant instructions, and the other service spacecraft are controlled based on the information between each other, and the information between each other includes the relative position and relative speed of the service spacecraft close to the fault spacecraft.

[0083] Preferably, the initial pipe connection position determination method is:

[0084] The mass center of the system composed of the multiple service spacecraft and the fault spacecraft coincides with the mass center of the fault spacecraft, and at this time, the multiple service spacecraft reach the initial pipe connection position.

[0085] Preferably, the initial pipe connection position satisfies

[0086] Wherein, m i represents the mass of the i th service spacecraft, the service spacecraft include a service spacecraft capable of receiving instructions and a service spacecraft incapable of receiving instructions, η i represents the relative position of the i th service spacecraft and the fault spacecraft, n represents the number of service spacecraft, and η0represents the coordinates of the mass center of the fault spacecraft.

[0087] The electromagnetic docking system coordinate system is preferably an orbit reference coordinate system, which is defined as a local vertical local horizontal (LVLH) coordinate system in the embodiment, the coordinate origin is the center of mass of the faulty spacecraft, the oy axis is perpendicular to the orbit plane of the electromagnetic docking and coincides with the negative direction of the orbital angular velocity of the faulty spacecraft, wherein the oz axis points to the center of the earth from the coordinate origin, and the ox axis satisfies the right-hand rule with the other two coordinate axes.

[0088] The mass, function and initial capture direction of each service spacecraft are preferably determined according to the capture background and task requirements, the initial system center of mass formed by the initial capture positions is substantially coincident with the center of mass of the faulty spacecraft, and the initial positions of the service spacecrafts can have slight deviations, so that the initial center of mass and the center of mass of the faulty spacecraft also have slight differences. The command trajectory of the electromagnetic docking and the command receiving and communication topology of each service spacecraft are determined according to the specific task, and in the controller, only the control force of the service spacecraft that can receive the command contains the command information, and the rest can only use the position and velocity information between the service spacecrafts. The electromagnetic force is:

[0089]

[0090] wherein μ C and μ T are the magnetic dipole moments of the two spacecrafts, μ0 is the vacuum permeability, and η is the relative position vector between the two spacecrafts. Since the magnetic moment of the faulty spacecraft is much larger than that of each service spacecraft, the electromagnetic force between the service spacecrafts can be assumed to be negligible.

[0091] The disturbance estimation value is obtained based on the relative information of the service spacecraft and the faulty spacecraft.

[0092] Preferably, the method for obtaining the disturbance estimation value is as follows:

[0093] The relative position of the service spacecraft and the faulty spacecraft is obtained.

[0094] The relative velocity of the service spacecraft and the faulty spacecraft is obtained based on the derivative of the relative position.

[0095] The disturbance estimation value is obtained by inputting the relative position and the relative velocity into the disturbance observer.

[0096] The relative position and the relative velocity are used as the relative information.

[0097] Preferably, since the electromagnetic docking is affected by external disturbances, in order to achieve high-precision docking, a disturbance observer is designed to observe external disturbances and compensate in the controller to improve the control precision of the electromagnetic capture.

[0098] A radial basis function (RBF) neural network disturbance observer is preferably designed, based on the relative position η i and relative velocity of the ith service spacecraft to the faulty spacecraft An action function is selected as a Gaussian function, and an adaptive learning rate is designed based on a sliding mode control to estimate the external disturbance of the system and obtain a disturbance estimation value

[0099]

[0100] h=[h j ] T ;

[0101] wherein represents a corresponding neural network weight matrix, represents an input of the neural network, h=[h j ] T represents an output of the Gaussian function, c j represents a center vector of the Gaussian function of the jth neuron, b j represents a standardization constant of the jth neuron.

[0102] The original moving direction is kept unchanged, and the relative position direction of the service spacecraft is adjusted to be consistent with the relative position direction of the service spacecraft capable of receiving the command, to obtain a rotated relative position.

[0103] Preferably, since the trajectory command is determined according to the relative position η k and mass m k of the service spacecraft capable of receiving the command, in order to keep the center of mass of the overall system unchanged, the other service spacecraft should make a similar motion trajectory in the direction of its own relative position η i according to its own mass m i , therefore, the relative position η i of each slave service spacecraft is adjusted to be consistent with the relative position direction of the service spacecraft capable of receiving the command by using the Rodrigues rotation formula, to obtain a rotated relative position: η ir =Rη i , wherein η ir represents the rotated relative position of the ith service spacecraft, and R represents a rotation matrix obtained by the Rodrigues rotation formula.

[0104] Based on the rotated relative position, the reference position trajectory and the disturbance estimation value, a control input is obtained.

[0105] Preferably, as shown in Figure 2 , the method for obtaining the control input is:

[0106] obtaining relative position error based on the rotational relative position and the reference position trajectory;

[0107] obtaining relative information error with other service spacecraft based on the relative position error and the communication topology relationship between the service spacecraft;

[0108] obtaining control force based on the relative information error;

[0109] obtaining control input based on the control force and the corresponding disturbance estimation value.

[0110] Preferably, the method for obtaining the relative position error is:

[0111] obtaining mass coefficient based on the mass of the service spacecraft and the mass of the service spacecraft capable of receiving instructions;

[0112] obtaining transformed relative position based on the rotational relative position and the corresponding mass coefficient;

[0113] obtaining relative position error based on the transformed relative position and the reference position trajectory.

[0114] Preferably, the mass of the ith service spacecraft m i is divided by the mass of the service spacecraft capable of receiving instructions m k to obtain the ith mass coefficient L i : L i = m k / m ir , the rotational relative position η i of the ith service spacecraft is multiplied by the corresponding mass coefficient L id to obtain the transformed relative position η id of the ith service spacecraft: η ir = η i × L id , the transformed relative position η d is used to obtain the relative position error e i : e i = η id - η d , and the problem is converted into a consensus problem of multiple servers.

[0115] Preferably, the relative information error z i with other service spacecraft is obtained based on the relative position error e ij and the communication topology relationship a i between the service spacecraft:

[0116]

[0117] wherein μ idenotes whether the ith service spacecraft can receive the reference position trajectory η d , if yes, μ i = 1, otherwise μ i = 0, n denotes the number of service spacecrafts, e j denotes the relative position error of the jth service spacecraft.

[0118] Preferably, the method for obtaining the control force is:

[0119] a sliding mode surface is obtained based on the derivative of the rotating relative position and the relative information error;

[0120] a control force after adjustment of direction is obtained based on the sliding mode surface, the derivative of the rotating relative position, the rotating relative position, the derivative of the relative information error and the relative information error.

[0121] Preferably, the sliding mode surface is obtained based on the derivative of the rotating relative position i and the relative information error z wherein f1(·) denotes a sliding mode surface function;

[0122] Secondly, a sliding mode reaching law is designed which can be obtained from the orbit dynamics equation i.e. is a function of u ir , and the control force after adjustment of direction is obtained by the derivative of the sliding mode surface with respect to time and the sliding mode reaching rate f4() denotes a control law in the direction after rotation.

[0123] Preferably, the control input ui: is obtained based on the control force u ir and the corresponding disturbance estimate

[0124] Based on the control input, each service spacecraft and the faulty spacecraft implement docking, and complete the takeover of the faulty spacecraft.

[0125] Based on the service spacecraft capable of receiving instructions, the control information is received and distributed to the actuators of all service spacecrafts, to realize the takeover control of the faulty spacecraft.

[0126] Preferably, after the electromagnetic docking is completed, each service spacecraft and the faulty spacecraft form an integral whole, and the control force and control torque are generated by the actuators of the service spacecrafts, to realize the attitude and orbit control of the faulty spacecraft.

[0127] In order to realize the attitude control and orbit control, the desired control force and control torque need to be obtained according to a certain control law, and then distributed to the actuators of each service spacecraft to generate actual control force and control torque.

[0128] Preferably, the control of the failed spacecraft is implemented, specifically including:

[0129] Based on the communication between each service spacecraft and the use of the actuator of each service spacecraft, using graph theory knowledge, each actuator is taken as a node, and the communication between each service spacecraft is abstracted as an edge connecting the nodes;

[0130] Based on each edge, a corresponding influence coefficient is given;

[0131] Based on the output torque, thruster life and energy consumption, a comprehensive factor is set for each node;

[0132] Based on the comprehensive factor, the control information and the influence coefficient, a torque distribution objective function containing constraints is constructed;

[0133] Based on the torque distribution objective function, the control force and the control torque on each actuator are obtained;

[0134] Based on the control force and the control torque, the control of the failed spacecraft is implemented.

[0135] Preferably, the comprehensive factor specifically includes:

[0136] Based on the accuracy of the output torque and the saturation phenomenon of the flywheel actuator, an ability factor is set;

[0137] Based on the thruster life, a life factor is set;

[0138] Based on the energy consumption, a performance index is set;

[0139] The ability factor, the life factor and the performance index jointly constitute the comprehensive factor.

[0140] Preferably, the design performance index primarily considers the actual output torque accuracy problem, and simultaneously considers the saturation phenomenon of the flywheel actuator, and based on this, the ability factor g ωi = f g1 (ω i ) is set, wherein ω i represents the angular velocity of the i-th flywheel, the ability factor is affected by the saturation degree of the flywheel, the smaller the ability factor corresponding to the flywheel with high saturation degree is, and vice versa, that is, f g1 (·) is a monotonically decreasing function.

[0141] Preferably, since the thruster needs to consume energy, too much energy consumed by a thruster may lose life prematurely, and therefore a life factor g ni = f g2 (n i ) is set for the thruster, wherein n iFor the number of times of using thrusters, the thruster with less use has higher use opportunity, i.e. f g2 (·) is also a monotonic decreasing function.

[0142] Preferably, based on the overall energy consumption, the performance index g εi = f g3 (ε i ), wherein ε i represents the remaining energy of the i th actuator, and for the thruster, it represents the proportion of the remaining fuel, and for the flywheel, it represents the proportion between the current angular momentum and the saturated angular momentum.

[0143] Preferably, the torque distribution target function is specifically:

[0144]

[0145] s.t.Ak i = τ d ;

[0146] A = [D ij ] ∈ R n×n ;

[0147] Wherein β 1i , β 2i and β 3i all represent weight coefficients, g ωi represents the ability factor, g ni represents the life factor, g εi represents the performance index, k i represents the output of the i th actuator, τ d represents the control information, D ij represents the influence coefficient, and n represents the number of service spacecraft.

[0148] Preferably, the torque distribution target function is solved based on the use of Lagrange multiplier method, and the control force and control torque distributed on each actuator are obtained, so as to realize the attitude and orbit control of the failed spacecraft.

[0149] Embodiment 3

[0150] As shown in Figure 3 , a cooperative electromagnetic takeover system for large-scale failed spacecraft includes a trajectory determination module, an initial position determination module, a disturbance estimation module, a direction adjustment module, a final docking module and a control module.

[0151] The trajectory determination module is used to determine the reference position trajectory based on the task demand;

[0152] The initial position determination module is used to release a plurality of service spacecraft to reach the initial takeover position;

[0153] The disturbance estimation module is used to obtain the corresponding disturbance estimate based on the relative information between the servicing spacecraft and the faulty spacecraft;

[0154] The orientation adjustment module is used to maintain the original direction of motion and adjust the relative position orientation of the service spacecraft so that it is consistent with the relative position orientation of the service spacecraft that can receive commands, thereby obtaining the rotational relative position.

[0155] The final docking module is used to obtain control inputs based on the relative rotation position, reference position trajectory, and disturbance estimate; based on the control inputs, each service spacecraft docks with the faulty spacecraft.

[0156] The control module is used to receive control information from service spacecraft capable of receiving commands and distribute it to the actuators of all service spacecraft to achieve control of the faulty spacecraft.

[0157] Preferably, the functions of each module in this embodiment correspond one-to-one with the above-described method, and will not be described in detail here.

[0158] Example 4

[0159] Based on the same inventive concept, the present invention also provides a computer device, including a processor, a communication interface, a memory and a communication bus, wherein the processor, the communication interface and the memory communicate with each other through the communication bus;

[0160] Memory, used to store computer programs;

[0161] When the processor executes a program stored in memory, it can implement a cooperative electromagnetic takeover method for a large-scale faulty spacecraft, as described in Embodiment 1 or 2.

[0162] like Figure 4 As shown, the electronic device may include a processor 41, a communications interface 42, a memory 43, and a communication bus 44, wherein the processor 41, the communications interface 42, and the memory 43 communicate with each other via the communication bus 44. The processor 41 can call logical instructions in the memory 43 to execute a cooperative electromagnetic takeover method for a large-scale faulty spacecraft as described in Embodiment 1 or 2.

[0163] Furthermore, the logic instructions in the aforementioned memory 43 can be implemented as software functional units and, when sold or used as independent products, can be stored in a computer-readable storage medium.

[0164] Based on such understanding, the technical solutions of the present application essentially or the parts of the prior art that contribute to the present application or the parts of the technical solutions can be embodied in the form of a software product, which is stored in a storage medium and includes a plurality of instructions for causing a computer device (which can be a personal computer, a server, or a network device, etc.) to execute all or part of the steps of the various embodiment methods of the present application.

[0165] The aforementioned storage medium includes a U disk, a mobile hard disk, a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disk, and various storage program code media.

[0166] Via the above technical solutions, the present application provides a cooperative electromagnetic takeover method and system for large-scale fault spacecraft, which uses electromagnetic force as the control force to effectively reduce the impact force of the contact type takeover and avoid causing secondary damage; and the electromagnetic force used to realize the takeover has the advantages of high fault tolerance, no rigid collision, no debris generation, no plume pollution, flexibility, wide application range, etc.; the present application only has one service spacecraft that can receive instructions, and the rest of the service spacecraft perform electromagnetic takeover according to the information between each other and keep the system centroid composed of each service spacecraft and the fault spacecraft unchanged, while reducing the memory requirement; after the docking is completed, the present application uses the actuators of each service spacecraft according to the designed torque distribution rules, so as to realize the accurate control of the attitude and orbit of the fault spacecraft.

[0167] The motor drive control algorithm is applied to the takeover control of large spacecraft, and the combined spacecraft after the takeover is controlled in attitude and orbit. First, the cooperative electromagnetic docking is used to connect the multiple service spacecraft and the fault spacecraft to form a combined spacecraft. There are a large number of nonlinear parts in the attitude and orbit model of the combined spacecraft, the nonlinear relationship of the controlled object is studied by using the motor drive control algorithm, the nonlinear model of the attitude and orbit coupling dynamics equation is established, the feedback compensation method is used to improve the response speed and stability of the motor driven actuator for the control command delay caused by the nonlinearity, the control accuracy of the combined spacecraft is improved due to the strong adaptability to the nonlinearity.

[0168] The embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts of each embodiment can be referred to each other. For the device disclosed in the embodiments, since it corresponds to the method disclosed in the embodiments, the description is relatively simple, and the related parts can be referred to the method part.

[0169] The foregoing description of the disclosed embodiments enables a person skilled in the art to make or use the application. Modifications of these embodiments will occur to persons of skill in the art, and, while certain modifications are discussed, it is desired to be protected in accordance with the spirit and scope of the application. Therefore, the application is not limited to the specific embodiments shown and described, but only by the scope of the appended claims, unless otherwise specified.

Claims

1. A collaborative electromagnetic takeover method for large-scale faulty spacecraft, characterized in that, include: Determine the reference location trajectory based on task requirements; Multiple service spacecraft were deployed to their initial takeover positions; The method for determining the initial pipe connection location is as follows: The centroid of the system consisting of multiple service spacecraft and the faulty spacecraft coincides with the centroid of the faulty spacecraft, at which point the multiple service spacecraft arrive at the initial takeover position. The corresponding disturbance estimate is obtained based on the relative information between the service spacecraft and the faulty spacecraft; The method for obtaining the disturbance estimate is as follows: Obtain the relative positions of the service spacecraft and the faulty spacecraft; The relative velocity between the service spacecraft and the faulty spacecraft is obtained by differentiating based on the relative positions. The disturbance estimate is obtained by inputting the relative position and the relative velocity to the disturbance observer. The relative position and the relative velocity serve as the relative information; Keeping the original direction of motion unchanged, adjust the relative position of the service spacecraft so that it is consistent with the relative position of the service spacecraft that can receive commands, and obtain the rotational relative position; Based on the relative rotation position, the reference position trajectory, and the disturbance estimate, the control input is obtained; The method for obtaining the control input is as follows: The relative position error is obtained based on the rotational relative position and the reference position trajectory; The relative information error with other service spacecraft is obtained based on the relative position error and the communication topology relationship between the service spacecraft. The control force is obtained based on the relative information error; The control input is obtained based on the control force and the corresponding disturbance estimate. Based on the control input, each of the service spacecraft and the faulty spacecraft docks; Based on the fact that the service spacecraft capable of receiving instructions receives control information and distributes it to the actuators of all service spacecraft, it can take over control of the faulty spacecraft. To achieve takeover control of the faulty spacecraft, specifically including: Based on the communication between the various service spacecraft and the usage of the actuators of the various service spacecraft, graph theory is used to treat each actuator as a node and the communication between the various service spacecraft as edges connecting the nodes. Each edge is assigned a corresponding influence coefficient. A comprehensive factor is set for each node based on output torque, thruster life, and energy consumption; Based on the comprehensive factor, the control information, and the influence coefficient, a constrained torque allocation objective function is constructed. Based on the torque distribution objective function, the control force and control torque on each actuator are obtained; The control force and the control torque are used to control the faulty spacecraft.

2. The collaborative electromagnetic takeover method for large-scale faulty spacecraft according to claim 1, characterized in that, The method for obtaining the relative position error is as follows: A mass coefficient is obtained based on the ratio of the mass of the service spacecraft to the mass of the service spacecraft capable of receiving commands. Based on the rotational relative position and the corresponding mass coefficient, the transformation relative position is obtained; The relative position error is obtained based on the transformed relative position and the reference position trajectory.

3. The collaborative electromagnetic takeover method for large-scale faulty spacecraft according to claim 1, characterized in that, The method for obtaining the control force is as follows: The sliding mode surface is obtained based on the derivative of the relative rotational position and the relative information error; The control force after adjusting the direction is obtained based on the sliding modal surface, the derivative of the rotational relative position, the rotational relative position, the derivative of the relative information error, and the relative information error.

4. The collaborative electromagnetic takeover method for large-scale faulty spacecraft according to claim 1, characterized in that, The comprehensive factors specifically include: Based on the accuracy of the output torque and the saturation phenomenon of the flywheel actuator, a capability factor is set; A life factor is set based on the thruster's lifespan; Based on the energy consumption, performance indicators are set; The comprehensive factor is composed of the capability factor, the lifespan factor, and the performance index.

5. A collaborative electromagnetic takeover method for large-scale faulty spacecraft according to claim 1, characterized in that, The specific objective function for torque distribution is: ； ； A =[ D ij ]∈ R n×n ； in, β 1i , β 2i and β 3i All represent weighting coefficients. Indicates ability factor, Represents the lifespan factor. Indicates performance metrics, k i Indicates the first i The output of each actuator Indicates control information, D ij Indicates the influence coefficient. n Indicates the number of spacecraft being served.

6. A cooperative electromagnetic takeover system for large-scale malfunctioning spacecraft, applied to the cooperative electromagnetic takeover method for large-scale malfunctioning spacecraft as described in any one of claims 1-5, characterized in that, include: The system includes a trajectory determination module, an initial position determination module, a disturbance estimation module, a direction adjustment module, a final docking module, and a control module. The trajectory determination module is used to determine the trajectory of the reference position based on the task requirements; The initial position determination module is used to release multiple service spacecraft to arrive at the initial takeover position; The disturbance estimation module is used to obtain the corresponding disturbance estimation value based on the relative information between the service spacecraft and the faulty spacecraft; The orientation adjustment module is used to maintain the original direction of motion and adjust the relative position orientation of the service spacecraft so that it is consistent with the relative position orientation of the service spacecraft that can receive commands, thereby obtaining a rotational relative position. The final docking module is used to obtain control input based on the relative rotation position, the reference position trajectory, and the disturbance estimate. Based on the control input, each of the service spacecraft and the faulty spacecraft docks; The control module is used to receive control information based on the service spacecraft capable of receiving instructions and distribute it to the actuators of all service spacecraft, thereby realizing the takeover control of the faulty spacecraft.

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