A multi-spacecraft attitude and orbit coupling full-drive convergence time controller generation method

Abstract

The application discloses a kind of multi-spacecraft attitude and orbit coupling full-drive convergence time controller generation method, applied to aerospace technology field, comprising: the full-drive characteristic of tracking error model is established multi-spacecraft attitude and orbit coupling error system full-drive model;Wherein, multi-spacecraft attitude and orbit coupling error system full-drive model is the pseudo-linear system model for the stability problem of multi-spacecraft attitude and orbit coupling error system;For multi-spacecraft attitude and orbit coupling error system full-drive model, based on full-drive system method, sliding mode control method and predetermined time control method, constructs full-drive convergence time controller, to adjust the convergence time of system based on full-drive convergence time controller;Full-drive convergence time controller is the controller designed based on predetermined time parameter and adjustable time parameter.The application proposes a kind of attitude and orbit coupling adjustable predetermined time method, so that multi-spacecraft can be maneuvered to the desired attitude and orbit within predetermined time, complete formation, improve the efficiency of formation.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, in particular to a multi-spacecraft attitude and orbit coupling full-drive convergence time controller generation method. BACKGROUND

[0002] In multi-spacecraft formation flight control, the research on attitude and orbit coupling control method is a key point. However, most of the existing methods can only guarantee the asymptotic stability of the control system, the convergence in a finite time or the convergence in a fixed time. The upper bound of the convergence time of the system cannot be predetermined in advance. Therefore, it is urgent to study the multi-spacecraft attitude and orbit coupling predetermined time control method. The existing predetermined time control method cannot further adjust the speed of time after the predetermined time upper bound.

[0003] Therefore, how to further adjust the convergence time is a technical problem that those skilled in the art urgently need to solve. SUMMARY

[0004] Therefore, the purpose of the present application is to provide a multi-spacecraft attitude and orbit coupling full-drive convergence time controller generation method, which solves the technical problem that the convergence time of the multi-spacecraft attitude and orbit coupling process in the prior art cannot be adjusted according to demand.

[0005] To solve the above technical problems, the present application provides a multi-spacecraft attitude and orbit coupling full-drive convergence time controller generation method, comprising:

[0006] establishing a tracking error model based on the attitude dynamics model and the attitude and orbit coupling mathematical model of the spacecraft;

[0007] establishing a full-drive model of the multi-spacecraft attitude and orbit coupling error system based on the full-drive characteristics of the tracking error model; wherein the full-drive model of the multi-spacecraft attitude and orbit coupling error system is a pseudo-linear system model for the stability problem of the multi-spacecraft attitude and orbit coupling error system;

[0008] designing a full-drive convergence time controller based on the full-drive system method, the sliding mode control method and the predetermined time control method for the full-drive model of the multi-spacecraft attitude and orbit coupling error system, to adjust the convergence time of the system based on the full-drive convergence time controller; wherein the full-drive convergence time controller is a controller designed based on predetermined time parameters and adjustable time parameters.

[0009] Optionally, establishing a tracking error model based on the attitude dynamics model and the attitude and orbit coupling mathematical model of the spacecraft, comprising:

[0010] establishing the attitude and orbit coupling mathematical model according to the attitude dynamics model of the spacecraft and the relative orbit dynamics;

[0011] determining a desired state, performing algebraic manipulation on the attitude dynamics model and the attitude-orbit coupled mathematical model based on the desired state, and establishing the tracking error model.

[0012] Optionally, the full-drive model of the multi-spacecraft attitude-orbit coupled error system is determined as ; wherein, is an attitude-orbit coupled error state variable, is a full-drive convergence time controller to be designed, is a total disturbance in the current system, represents a second-order derivative of the attitude-orbit coupled error state variable.

[0013] Optionally, the full-drive convergence time controller is designed based on a full-drive system method, a sliding mode control method, and a predetermined time control method for the full-drive model of the multi-spacecraft attitude-orbit coupled error system, and includes:

[0014] constructing a consensus error based on information of adjacent neighbor spacecrafts of a current spacecraft;

[0015] constructing a consensus error function based on the consensus error; wherein the consensus error function is used to guarantee a predetermined time convergence characteristic;

[0016] constructing a sliding mode surface based on the consensus error and the consensus error function;

[0017] designing the full-drive convergence time controller based on the sliding mode surface, the consensus error, a predetermined time parameter, and an adjustable time parameter.

[0018] Optionally, the full-drive convergence time controller is ; wherein, and are parameters determined based on a topology of the multi-spacecraft. is a control parameter, is an estimated value of a total disturbance of the system, is a sliding mode surface, is a consensus error function, is a sliding mode surface predetermined time function, is a full-drive convergence time controller of the jth spacecraft;

[0019] wherein, , is a consensus error, , , and is a predetermined time parameter, and is an adjustable time parameter.

[0020] Optionally, the spacecraft is a rigid spacecraft with external disturbance and actuator failure.

[0021] Optionally, before the tracking error model is established based on the attitude dynamics model and the attitude-orbit coupling mathematical model, the method further comprises:

[0022] The multi-spacecraft scenario is determined to be a formation flight task composed of one virtual leader spacecraft and n follower spacecrafts.

[0023] The attitude dynamics model corresponding to the multi-spacecraft scenario is determined to be ; wherein, is a modified Rodrigues parameter, is an angular velocity, is a spacecraft moment of inertia, is a control torque, is a disturbance torque, , , is a physically meaningless parameter, denotes the first derivative of the modified Rodrigues parameter, denotes the second derivative of the modified Rodrigues parameter.

[0024] The present application also provides a convergence time adjustment method, comprising:

[0025] Obtaining multi-spacecraft input parameters;

[0026] Based on the multi-spacecraft input parameters, the convergence time of the attitude-orbit coupling is controlled by using a full-drive convergence time controller to obtain target state variables; wherein the full-drive convergence time controller is obtained based on the above multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method.

[0027] Optionally, the multi-spacecraft input parameters include a sliding mode surface parameter and a total disturbance of the system determined based on an observer.

[0028] Optionally, after the convergence time of the attitude-orbit coupling is controlled by using a convergence time controller based on the multi-spacecraft input parameters to obtain target state variables, the method further comprises:

[0029] The target state variables are sent to an actuator failure module to make the actuator failure module generate an attitude change signal, and the state of the spacecraft is adjusted based on the attitude change signal.

[0030] The present application also provides a multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation implementation device, comprising:

[0031] The tracking error model determining module is configured to establish a tracking error model based on an attitude dynamics model and an orbit-attitude coupling mathematical model of the spacecraft.

[0032] The multi-spacecraft orbit-attitude coupling error system full-drive model establishing module is configured to establish a multi-spacecraft orbit-attitude coupling error system full-drive model based on full-drive characteristics of the tracking error model.

[0033] The full-drive convergence time controller implementing module is configured to design a full-drive convergence time controller based on a full-drive system method, a sliding mode control method and a predetermined time control method for the multi-spacecraft orbit-attitude coupling error system full-drive model, so as to adjust the convergence time of the system based on the full-drive convergence time controller.

[0034] The present application further provides a convergence time adjusting method, comprising:

[0035] The input parameter acquiring module is configured to acquire multi-spacecraft input parameters.

[0036] The convergence time adjusting module is configured to control the convergence time of the orbit-attitude coupling by using the full-drive convergence time controller based on the multi-spacecraft input parameters, so as to obtain a target state variable.

[0037] The present application further provides an electronic device, comprising:

[0038] The memory is configured to store a computer program.

[0039] The processor is configured to execute the computer program to implement the steps of the multi-spacecraft orbit-attitude coupling full-drive convergence time controller generating method and the convergence time adjusting method.

[0040] The present application further provides a computer readable storage medium for storing a computer program, wherein the computer program is executed by a processor to implement the multi-spacecraft orbit-attitude coupling full-drive convergence time controller generating method and the convergence time adjusting method.

[0041] The present application further provides a computer program product comprising a computer program / instruction, wherein the computer program / instruction is executed by a processor to implement the steps of the multi-spacecraft orbit-attitude coupling full-drive convergence time controller generating method and the convergence time adjusting method.

[0042] It can be seen that the application establishes a tracking error model based on an attitude dynamics model and an attitude-orbit coupling mathematical model of a spacecraft; establishes a full-drive model of a multi-spacecraft attitude-orbit coupling error system based on full-drive characteristics of the tracking error model; wherein the full-drive model of the multi-spacecraft attitude-orbit coupling error system is a pseudo-linear system model for the stability problem of the multi-spacecraft attitude-orbit coupling error system; a full-drive convergence time controller is designed based on a full-drive system method, a sliding mode control method and a predetermined time control method for the full-drive model of the multi-spacecraft attitude-orbit coupling error system, so as to adjust the convergence time of the system based on the full-drive convergence time controller; wherein the full-drive convergence time controller is a controller designed based on a predetermined time parameter and an adjustable time parameter. The beneficial effects of the application are that, compared with the current convergence time which cannot be further adjusted, the full-drive convergence time controller based on the predetermined time parameter and the adjustable time parameter is designed to realize that the multi-spacecraft can be maneuvered to the expected attitude and orbit within the predetermined time, complete the formation and improve the efficiency of the formation.

[0043] In addition, the application also provides a convergence time adjustment method, which also has the above beneficial effects. BRIEF DESCRIPTION OF DRAWINGS

[0044] In order to more clearly illustrate the technical solutions in the embodiments of the application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or the prior art description. Obviously, the drawings in the following description are only the embodiments of the application, and for those skilled in the art, other drawings can also be obtained without creative labor based on the provided drawings.

[0045] Figure 1 A flow chart of a multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method provided by the embodiment of the application;

[0046] Figure 2 A flow chart of a multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method provided by the embodiment of the application;

[0047] Figure 3 A control system block diagram provided by the embodiment of the application;

[0048] Figure 4 A communication topology schematic diagram provided by the embodiment of the application;

[0049] Figure 5 An actual formation state provided by the embodiment of the application;

[0050] Figure 6 An attitude variable and attitude angular velocity schematic diagram provided by the embodiment of the application;

[0051] Figure 7A position variable and a speed variable schematic diagram provided for the embodiment of the present application;

[0052] Figure 8 A control torque and control force curve under the action of a full-drive convergence time controller provided for the embodiment of the present application;

[0053] Figure 9 A flowchart of a convergence time adjustment method provided for the embodiment of the present application. DETAILED DESCRIPTION

[0054] To make the objectives, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by a person of ordinary skill in the art without creative work fall within the protection scope of the present application.

[0055] Please refer to Figure 1 , Figure 1 A flowchart of a multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method provided for the embodiment of the present application. The method can include:

[0056] S101, establishing a tracking error model based on an attitude dynamics model and an attitude-orbit coupling mathematical model of a spacecraft.

[0057] The execution subject of the embodiment is an electronic device, which can be specifically a server, a portable terminal or other forms. The spacecraft in the embodiment is the i th spacecraft in the multi-spacecraft. The attitude dynamics model in the embodiment is a model reasoned out according to an existing model, and the application scenario is a formation flight task composed of 1 virtual leader spacecraft and n follower spacecrafts. The embodiment establishes a tracking error model by inputting a given expectation.

[0058] It needs to be further explained that, based on any of the above embodiments, establishing a tracking error model based on an attitude dynamics model and an attitude-orbit coupling mathematical model of a spacecraft can include:

[0059] S1011, establishing an attitude-orbit coupling mathematical model according to the attitude dynamics model and the relative orbit dynamics of the spacecraft.

[0060] The embodiment establishes an attitude-orbit coupling mathematical model according to the attitude dynamics and the relative orbit dynamics of the spacecraft in the presence of external disturbance and actuator failure.

[0061] S1012, determine the desired state, based on the desired state, algebraic change is made to the attitude dynamics model and the attitude orbit coupling mathematical model, and a tracking error model is established.

[0062] This embodiment establishes a tracking error model by inputting a given expectation (which is obtained by algebraic transformation of the attitude dynamics and attitude orbit coupling mathematical model). This embodiment improves the accuracy of establishing a tracking error model by giving a specific process of establishing a tracking error model.

[0063] It should be further explained that, based on any of the above embodiments, before establishing the tracking error model based on the attitude dynamics model and the attitude orbit coupling mathematical model of the spacecraft, it can also include: determining that the multi-spacecraft scenario is a formation flight task composed of one virtual leader spacecraft and n follower spacecrafts; determining that the attitude dynamics model corresponding to the multi-spacecraft scenario is ; wherein, is the modified Rodrigues parameter, is the angular velocity, is the moment of inertia of the spacecraft, is the control torque, is the disturbance torque, , , is a physically meaningless parameter. This embodiment gives a specific attitude dynamics model corresponding to the formation flight task composed of one virtual leader spacecraft and n follower spacecrafts.

[0064] S102, based on the full drive characteristics of the tracking error model, a multi-spacecraft attitude orbit coupling error system full drive model is established; wherein the multi-spacecraft attitude orbit coupling error system full drive model is a pseudo-linear system model for the stability problem of the multi-spacecraft attitude orbit coupling error system.

[0065] This embodiment establishes a multi-spacecraft attitude orbit coupling error system full drive model based on the full drive characteristics of the tracking error model, which can simplify the multi-spacecraft attitude orbit coupling control problem into the stability problem of the multi-spacecraft attitude orbit coupling error system. At the same time, since the system model has been transformed into a kind of pseudo-linear system model, the design of the controller will also be simpler. It should be further explained that the multi-spacecraft attitude orbit coupling error system full drive model is ; wherein, is the attitude orbit coupling error state variable, is the full drive convergence time controller to be designed, is the total disturbance in the current system, Second derivative of the state variable representing the attitude and orbit coupling error. In this embodiment, a dot above each letter represents a first derivative, and two dots represent a second derivative. The system model in this embodiment (i.e., the full drive model of the multi-spacecraft attitude and orbit coupling error system) has been converted into a class of pseudo-linear system models, and the controller design based on this will also be simpler.

[0066] S103, based on the full drive system method, the sliding mode control method and the predetermined time control method, a full drive convergence time controller is designed for the full drive model of the multi-spacecraft attitude and orbit coupling error system, so as to adjust the convergence time of the system based on the full drive convergence time controller; wherein the full drive convergence time controller is a controller designed based on the predetermined time parameter and the adjustable time parameter.

[0067] The full drive convergence time controller in this embodiment mainly adjusts the convergence time of the system through the adjustable time parameter. Under the action of the full drive convergence time controller in this embodiment, each state of the system can converge within a predetermined time, that is, the multi-spacecraft can complete the formation task within a predetermined time in the presence of external disturbances and internal actuator faults. The full drive convergence time controller in this embodiment is a full drive controller designed based on the full drive system theory, and also combines the predetermined time control method, so it is a full drive convergence time controller.

[0068] It should be further pointed out that based on any of the above embodiments, the full drive convergence time controller is designed based on the full drive system method, the sliding mode control method and the predetermined time control method for the full drive model of the multi-spacecraft attitude and orbit coupling error system, which can include:

[0069] S1031, constructing a consensus error based on the information of the adjacent neighbor spacecraft of the current spacecraft.

[0070] One key point of the multi-spacecraft formation controller in this embodiment is to know the information of the neighbor spacecraft for cooperative control, so as to construct a consensus error. The consensus error is the result of the difference between the state information of the neighbor spacecraft and the state information of the spacecraft itself, and represents the error information between the spacecraft and the neighbor spacecraft.

[0071] S1032, designing a consensus error function based on the consensus error; wherein the consensus error function is used to ensure the predetermined time convergence characteristic.

[0072] The consensus error function in this embodiment is a function of the consensus error, which is used to ensure the predetermined time convergence characteristic.

[0073] S1033, designing a sliding surface based on the consensus error and the consensus error function.

[0074] S1034, designing the full-drive convergence time controller based on the sliding surface, the consistency error, the predetermined time parameter and the adjustable time parameter.

[0075] In this embodiment, the sliding surface predetermined time parameter can be set through the sliding surface, the consistency error predetermined time parameter can be set according to the consistency error, and the convergence time of the system can be further adjusted through the adjustable time parameter. This embodiment gives a specific method for designing the full-drive convergence time controller, and improves the accuracy of the full-drive convergence time controller design.

[0076] It should be further explained that, based on any of the above embodiments, the full-drive convergence time controller is ; wherein, and are parameters determined based on the topology of the multi-spacecraft. is a control parameter, is an estimated value of the total disturbance of the system, is a sliding surface, is the consistency error function, is the sliding surface predetermined time function, is the full-drive convergence time controller of the jth spacecraft;

[0077] wherein, , is the consistency error, , , and are predetermined time parameters, and are adjustable time parameters. Under the action of the full-drive convergence time controller in this embodiment, the states of the system can converge within the predetermined time , that is, the multi-spacecraft can complete the formation task within the predetermined time in the presence of external disturbances and internal actuator faults.

[0078] The multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method provided by the embodiment of the application can comprise: establishing a tracking error model based on an attitude dynamics model and an attitude-orbit coupling mathematical model of a spacecraft; establishing a full-drive model of a multi-spacecraft attitude-orbit coupling error system based on full-drive characteristics of the tracking error model; wherein the full-drive model of the multi-spacecraft attitude-orbit coupling error system is a pseudo-linear system model for the stability problem of the multi-spacecraft attitude-orbit coupling error system; a full-drive convergence time controller is designed based on a full-drive system method, a sliding mode control method and a predetermined time control method for the full-drive model of the multi-spacecraft attitude-orbit coupling error system, so as to adjust the convergence time of the system based on the full-drive convergence time controller; wherein the full-drive convergence time controller is a controller designed based on a predetermined time parameter and an adjustable time parameter. The controller structure designed by the full-drive system theory is simpler than the existing multi-spacecraft formation flight attitude-orbit coupling controller. In addition, after the upper limit of the convergence time of the system is predetermined in most existing predetermined time controls, the convergence time of the system cannot be further adjusted. Because the adjustable time parameter exists in the controller in the application, adjusting the adjustable time parameter can change the characteristics of the system, so that the adjustable full-drive convergence time controller can further accelerate the convergence of the system under the premise of the predetermined upper limit of the convergence time.

[0079] In order to make the application more convenient to understand, please refer to Figure 2 , Figure 2 The flowchart example of the multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method provided by the embodiment of the application can comprise:

[0080] S201: Constructing a full-drive model of a multi-spacecraft attitude-orbit coupling error system.

[0081] Consider a formation flight task composed of one virtual leader spacecraft and n follower spacecrafts. The attitude dynamics model of the i th spacecraft is:

[0082] (1);

[0083] wherein, is a modified Rodrigues parameter, is an angular velocity, is the rotational inertia of the spacecraft, is a control torque, is a disturbance torque. , , The form of (no physical meaning) is:

[0084] .

[0085] For the vector , denotes an anti-symmetric matrix, and has the form

[0086] .

[0087] The relative orbital dynamics model of the ith spacecraft is

[0088] (2);

[0089] where, is the position information, is the control force, is the disturbance force, is the mass of the spacecraft. , and the rotation matrix has the form

[0090] .

[0091] In the above formula, C and N are the products of algebraic transformation and have no physical meaning, the number in the lower right corner represents the row number, is the earth gravity parameter, and are the distances from the center of the earth to the virtual leader spacecraft and the ith follower spacecraft, respectively, denote the perigee angle, the true anomaly angle, the orbital inclination angle, and the ascending node right ascension of the virtual leader spacecraft, respectively. The expression is as follows:

[0092] .

[0093] Let the attitude-orbit coupled state variable of the spacecraft be , the coupled controller be , and the coupled disturbance be . Considering the actuator fault of the spacecraft, the actuator fault is modeled as , denotes the ideal control input, where is the efficiency factor, is the drift fault. Let the error state be , is the desired state, and the letter above e represents error and the letter above d represents desired. Thus, by combining formula (1) and formula (2), the attitude-orbit coupled control error system model of the ith spacecraft is obtained as follows:

[0094] (3);

[0095] In the formula,

[0096] ; where C is the state matrix, N is the known part of the system, and H is the input matrix.

[0097] The total disturbance of the system is estimated by the observer and compensated into the system. The expression is:

[0098] .

[0099] where, denotes the 6-dimensional identity matrix.

[0100] .

[0101] Obviously, the control matrix is invertible, so by design:

[0102] (4) ;

[0103] Substituting (4) into (3), the system model (3) can be transformed into a full-drive model of the multi-spacecraft attitude-orbit coupling error system, i.e.:

[0104] (5) ;

[0105] Therefore, the core objective of the present application is to design a controller so that the system (5) can be stabilized within a predetermined time.

[0106] S202: Design a full-drive convergence time controller based on the full-drive model of the multi-spacecraft attitude-orbit coupling error system.

[0107] One key point of the multi-spacecraft formation controller is to know the information of the neighbor spacecraft for collaborative control. Therefore, first, a consensus error is constructed:

[0108] (6) ;

[0109] denotes the communication with the neighbor spacecraft, 1 for communication and 0 for no communication; denotes the error state of the jth spacecraft. According to equation (6), a sliding surface is designed. Where, is a function of the consensus error, which is used to guarantee the predetermined time convergence property, and has the following form:

[0110] .

[0111] In the formula, . , is a parameter, An adjustable time parameter for adjusting the convergence time.

[0112] Therefore, the main controller can be designed as follows:

[0113] (7).

[0114] in, and The parameters are known after the communication topology of the multiple spacecraft is determined. These are control parameters. This is an estimate of the total system disturbance. For sliding surface, , These are the consistency error predetermined time parameter and the sliding surface predetermined time parameter, respectively, and their specific forms are as follows:

[0115] ;

[0116] In the above formula, This is to address the consistency error in multi-spacecraft formation systems. , . and These are the scheduled time parameters; and This is an adjustable time parameter.

[0117] S203: The effectiveness of the all-drive convergence time controller is verified by simulating the attitude and orbit coupling control scenario of one virtual lead spacecraft and three follower spacecraft.

[0118] The simulation assumes one virtual lead spacecraft and three follower spacecraft. For easier understanding, please refer to... Figure 3 , Figure 3 This invention provides a control system block diagram according to an embodiment of the invention. The module for designing the all-drive convergence time controller in this invention is... Figure 3 The gray, yellow, and blue modules in the text Figure 3 The adjustable predetermined time attitude-orbit coupling controller mentioned above is the full-drive convergence time controller. This controller is designed primarily based on consistency error, observer output parameters, and sliding surface. First, the attitude-orbit coupling mathematical model of the i-th spacecraft is formed from the spacecraft's attitude dynamics and orbital dynamics. The attitude and orbital state variables output from this mathematical model are fed back to the input of the control system. The difference between this input and the desired state is used to obtain the error state. Based on the error state, a full-drive system model for spacecraft attitude-orbit coupling control is established. Uncertainties and external disturbances in the system are considered as the total disturbance, which is estimated by the observer (the total disturbance is composed of uncertainties and external disturbances in the system). Then, the full-drive convergence time controller is designed and applied to the spacecraft system.Figure 4 A communication topology diagram is provided for the embodiment of the present application. The virtual leader spacecraft is marked as '0', and '1', '2' and '3' represent the follower spacecrafts respectively.

[0119] The orbit parameters of the virtual leader spacecraft are as follows:

[0120] .

[0121] The simulation parameters of each follower spacecraft are shown in Table 1, which is a spacecraft formation parameter table provided by the present application.

[0122] Table 1 Spacecraft formation parameter table

[0123]

[0124] Suppose the disturbance torque acting on the spacecraft is:

[0125] .

[0126] The controller parameters are selected as:

[0127] .

[0128] The final simulation results are shown in Figures 5-8 . Figure 5 A practical formation state is provided for the embodiment of the present application, which shows the actual formation state of the spacecraft in the relative orbit coordinate system at a simulation time of 500 seconds. The round dots in the figure indicate the initial positions of the three follower spacecrafts, and the squares indicate the final positions. Thus, it can be intuitively presented that the three follower spacecrafts can fly to the circular orbit centered on the virtual leader spacecraft from different positions in space and be uniformly distributed on the orbit. Figure 6 An attitude variable and attitude angular velocity diagram is provided for the embodiment of the present application, through Figure 6 It can be seen that each state converges within a predetermined time of 60 seconds, and because of the effect of the control parameter Figure 7 A position variable and velocity variable diagram is provided for the embodiment of the present application, through Figure 7 It can be seen that each state converges within a predetermined time of 60 seconds, and because of the effect of the control parameter Figure 6 and Figure 7 are respectively the state variable response curves of the attitude and orbit of the three follower spacecrafts under the action of the full-drive convergence time controller. It can be seen that each state converges within a predetermined time of 60 seconds, and because of the effect of the adjustable time parameter , the convergence is faster. Figure 8 A control torque and control force curve under the action of the full-drive convergence time controller is provided for the embodiment of the present application, Figure 8The left side is the control torque curve, and the right side is the control force curve. Figure 8 The control torque and control force values at 0 time are shown, which reflect the maximum control torque and control force required by the controller, and the control torque and control force values now basically conform to actual application.

[0129] For ease of understanding, please refer to Figure 9 , Figure 9 A flowchart of a convergence time adjustment method provided by the embodiment of the application, and can specifically include:

[0130] S301, obtaining multi-spacecraft input parameters.

[0131] The multi-spacecraft input parameters in the embodiment are input parameters required by the full-drive convergence time controller under the current multi-spacecraft scene, and the multi-spacecraft in the embodiment can include rigid spacecraft. For example, the multi-spacecraft input parameters in the embodiment can include a sliding mode surface parameter and a system total disturbance determined based on an observer.

[0132] S302, controlling the convergence time of the attitude-orbit coupling based on the multi-spacecraft input parameters by using the full-drive convergence time controller to obtain target state variables.

[0133] The full-drive convergence time controller in the embodiment is a controller obtained based on the multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method.

[0134] It needs to be further explained that after the target state variables are obtained by controlling the convergence time of the attitude-orbit coupling based on the multi-spacecraft input parameters by using the full-drive convergence time controller, the method can further include: sending the target state variables to an actuator fault module, so that the actuator fault module generates an attitude change signal and adjusts the state of the spacecraft based on the attitude change signal. The embodiment can complete the formation task within a predetermined time through the attitude change signal.

[0135] Each embodiment in the specification is described in a progressive manner, and each embodiment focuses on the difference from other embodiments. The same or similar parts of each embodiment can be referred to each other. For the device disclosed by the embodiment, since it corresponds to the method disclosed by the embodiment, the description is relatively simple, and the related parts can be referred to the method part.

[0136] Those skilled in the art will further realize that the mere conception of the examples described herein is not inducing the patentable subject matter recited in each of the claims. Consequently, in light of the foregoing disclosure, various modifying and changing principles are intended to be included within the scope of the inventive subject matter described herein, as those skilled in the art realize. Therefore, the above description is intended to be illustrative rather than restrictive.

[0137] Finally, it is also necessary to point out that, in this paper, the relationship such as first and second belongs to only distinguish one entity or operation from another entity or operation, and does not necessarily require or imply any such actual relationship or order between the entities or operations. Moreover, the term "include", "contain" or other any variant is intended to cover non-exclusive inclusion, so that the process, method, article or equipment including a series of elements not only includes those elements, but also includes other elements not explicitly listed, or includes the elements inherent in such process, method, article or equipment.

[0138] The above is a detailed introduction to the method for generating a multi-spacecraft attitude and orbit coupling full-drive convergence time controller provided by the present application, and the principle and implementation mode of the present application are described by applying specific examples in this paper. The above example is only used to help understand the method of the present application and its core idea; at the same time, for those skilled in the art, according to the idea of the present application, the specific implementation mode and application range will be changed, and the above description should not be understood as a limitation of the present application.

Claims

1. A method for generating a multi-spacecraft attitude-orbit coupled full-actuated convergence time controller, characterized in that, include: A tracking error model is established based on the spacecraft's attitude dynamics model and attitude-orbit coupling mathematical model; establish a full-drive model of a multi-spacecraft attitude-orbit coupled error system based on full-drive characteristics of the tracking error model; wherein the full-drive model of the multi-spacecraft attitude-orbit coupled error system is a pseudo-linear system model for a stability problem of the multi-spacecraft attitude-orbit coupled error system; and determine that the full-drive model of the multi-spacecraft attitude-orbit coupled error system is ; wherein, is an attitude-orbit coupled error state variable, is a full-drive convergence time controller to be designed, is a total disturbance in a current system, represents a second-order derivative of the attitude-orbit coupled error state variable; For the full-drive model of the multi-spacecraft attitude-orbit coupling error system, based on the full-drive system method, sliding mode control method, and predetermined time control method, a full-drive convergence time controller is designed to adjust the convergence time of the system; wherein, the full-drive convergence time controller is a controller designed based on predetermined time parameters and adjustable time parameters. Specifically, for the multi-spacecraft attitude-orbit coupling error system full-drive model, based on the full-drive system method, sliding mode control method, and predetermined time control method, a full-drive convergence time controller is designed, including: A consistency error is constructed based on information from the neighboring spacecraft of the current spacecraft; A consistency error function is constructed based on the aforementioned consistency error; wherein, the consistency error function is used to guarantee convergence characteristics within a predetermined time. A sliding surface is constructed based on the consistency error and the consistency error function; The all-drive convergence time controller is designed based on the sliding surface, the consistency error, the predetermined time parameter, and the adjustable time parameter; Wherein, the full-drive convergence time controller is ; wherein, and is a parameter determined based on the topology of multiple spacecrafts, is a control parameter, is a sliding surface, is a consensus error function, is a sliding surface predetermined time function, is a full-drive convergence time controller of the jth spacecraft, and n is the number of follower spacecrafts, denotes the first derivative of the consensus error function; in, , For consistency error, , , and For the predetermined time parameter, and This is an adjustable time parameter. These are meaningless parameters.

2. The method for generating a multi-spacecraft attitude-orbit coupling all-drive convergence time controller according to claim 1, characterized in that, A tracking error model is established based on the spacecraft's attitude dynamics model and attitude-orbit coupling mathematical model, including: The attitude-orbit coupling mathematical model is established based on the spacecraft's attitude dynamics model and relative orbital dynamics. Determine the desired state, and based on the desired state, perform algebraic transformations on the attitude dynamics model and the attitude-track coupling mathematical model to establish the tracking error model.

3. The method for generating a multi-spacecraft attitude-orbit coupling all-drive convergence time controller according to claim 1, characterized in that, The spacecraft in question is one that is subject to external interference and actuator malfunction.

4. The method for generating a multi-spacecraft attitude-orbit coupling all-drive convergence time controller according to claim 1, characterized in that, Before establishing the tracking error model based on the spacecraft's attitude dynamics model and attitude-orbit coupling mathematical model, the following steps are also included: The multi-spacecraft scenario is defined as a formation flight mission consisting of one virtual lead spacecraft and n follower spacecraft; The attitude dynamics model corresponding to the multi-spacecraft scenario is determined as follows: ;in, For modified Rodriguez parameters, Angular velocity, For the spacecraft's rotational inertia, To control the torque, For the disturbance torque, , , Parameters with no physical meaning The first derivative of the modified Rodriguez parameter is given. denoted by .

5. A convergence time adjustment method, characterized in that, include: Acquire input parameters from multiple spacecraft; Based on the input parameters of the multiple spacecraft, the convergence time of attitude-orbit coupling is controlled by the full-drive convergence time controller to obtain the target state variable; wherein, the full-drive convergence time controller is a controller obtained based on the multi-spacecraft attitude-orbit coupling full-drive convergence time controller generation method according to any one of claims 1 to 4.

6. The convergence time adjustment method according to claim 5, characterized in that, The multi-spacecraft input parameters include sliding surface parameters and total system perturbation determined by the observer.

7. The convergence time adjustment method according to claim 5, characterized in that, After obtaining the target state variables by controlling the convergence time of attitude-orbit coupling using the all-drive convergence time controller based on the multi-spacecraft input parameters, the process further includes: The target state variable is sent to the actuator fault module so that the actuator fault module generates an attitude change signal and adjusts the spacecraft's state based on the attitude change signal.

Images