Modeling simulation method and device for on-orbit deployment of annular reflector antenna

Abstract

The invention discloses a modeling simulation method and device for on-orbit unfolding of an annular reflector antenna, and belongs to the field of spacecraft attitude control. The method comprises the following steps: calculating a geometric displacement curve of each discretized equivalent rigid body of an annular truss antenna according to a geometric type center change curve of a satellite reflector antenna in a ground unfolding process; according to the geometric displacement curve and the geometric installation relation of the reflector antenna, establishing a motion differential equation of the translational motion of the multiple equivalent rigid bodies of the reflector antenna; and performing numerical integration on the motion differential equation according to a fourth-order Runge-Kutta method, and determining a simulation model for representing satellite attitude motion change in the reflector antenna unfolding process. According to the method, the simulation complexity of the on-orbit expansion process of the large satellite can be greatly reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft attitude control, and in particular to a modeling and simulation method and device for on-orbit deployment of a ring reflector antenna. Background Art

[0002] Geosynchronous synthetic aperture radar satellites typically utilize large, trussed, mesh parabolic antennas for microwave imaging. Because the antenna's aperture is significantly larger than the launch vehicle's fairing envelope, it must be stowed at launch and deployed after entering the geosynchronous orbit. During in-orbit deployment of the reflector antenna, the satellite's dimensions, center of mass, and moment of inertia undergo drastic changes, significantly impacting the satellite's control system. To analyze the impact of these changes, simulations are conducted based on the reflector deployment motion obtained from ground testing, combined with the satellite's attitude dynamics and kinematic modeling.

[0003] In related technologies, the disturbance of satellite attitude is usually analyzed based on the force, torque and stress data of the ring truss mesh parabolic antenna in the ground deployment test. However, this method has a huge and complex computational load, which makes the method too costly and has certain processing difficulties in practical application.

[0004] Based on this, there is an urgent need for a modeling and simulation method and device for the on-orbit deployment of a ring reflector antenna to solve the above technical problems. Summary of the Invention

[0005] This invention provides a modeling and simulation method and device for the on-orbit deployment of a ring reflector antenna, which can greatly reduce the simulation complexity of the on-orbit deployment process of large satellites. The technical solution is as follows:

[0006] In one aspect, a modeling and simulation method for on-orbit deployment of a ring reflector antenna is provided, the method comprising:

[0007] According to the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground, the geometric displacement curves of the discretized equivalent rigid bodies of the ring truss antenna are calculated.

[0008] Establishing a differential equation of motion for the multi-equivalent rigid body translational motion of the reflector antenna based on the geometric displacement curve and the geometric installation relationship of the reflector antenna;

[0009] The motion differential equation is numerically integrated according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the deployment of the reflector antenna.

[0010] In another aspect, a modeling and simulation device for on-orbit deployment of a ring reflector antenna is provided, the device comprising:

[0011] A calculation module is used to calculate the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna according to the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground;

[0012] A modeling module, configured to establish a motion differential equation of a multi-equivalent rigid body translational motion of the reflector antenna based on the geometric displacement curve and the geometric installation relationship of the reflector antenna;

[0013] The determination module is used to perform numerical integration on the motion differential equation according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the deployment of the reflector antenna.

[0014] On the other hand, a computer device is provided, which includes a memory and a processor, wherein the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to implement the steps of the above-mentioned modeling and simulation method for on-orbit deployment of a ring reflector antenna.

[0015] On the other hand, a computer-readable storage medium is provided, wherein the storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the modeling and simulation method for the on-orbit deployment of the annular reflector antenna are implemented.

[0016] On the other hand, a computer program product is provided, comprising a computer program, which, when executed by a processor, implements the steps of the above-mentioned modeling and simulation method for on-orbit deployment of a ring reflector antenna.

[0017] The technical solution provided by the present invention can at least bring about the following beneficial effects: first, based on the geometric centroid change curve of the satellite reflector antenna during ground deployment, the geometric displacement curves of each discretized equivalent rigid body of the annular truss antenna are calculated; then, based on the geometric displacement curves of each discretized equivalent rigid body of the annular truss antenna and the geometric relationship of the installation of the annular truss antenna on the entire satellite, the attitude dynamics and kinematic equations considering the translational motion of multiple rigid bodies are numerically integrated to obtain the satellite attitude motion changes during the reflector antenna deployment process. The calculation process adopted by this method is simple and effective, and has strong practicality. At the same time, it can be extended to the disturbance analysis of other motion accessories on the satellite attitude, and it also has reference significance for the simulation analysis of the in-orbit deployment process of other large satellites. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0019] Figure 1 This is a flow chart of a modeling and simulation method for on-orbit deployment of a ring reflector antenna provided by one embodiment of the present invention;

[0020] Figure 2 Schematic diagram of the geometric configuration and coordinate system of a large annular reflector antenna and a satellite center body provided by one embodiment of the present invention;

[0021] Figure 3 1 is a schematic diagram of a geometric centroid displacement curve along the X, Y, and Z axes in the local coordinate system of the antenna during ground deployment of a reflector antenna provided by one embodiment of the present invention;

[0022] Figure 4 1 is a schematic diagram of a geometric centroid displacement curve along the X and Y axes of an equivalent rigid body in the local coordinate system of an antenna provided by an embodiment of the present invention;

[0023] Figure 5 1. A schematic diagram of a curve showing changes in satellite roll, pitch, and yaw attitude angles during antenna deployment according to an embodiment of the present invention;

[0024] Figure 6 Schematic diagram of the angular velocity change curve of the satellite along the X-axis, Y-axis, and Z-axis of the satellite during the antenna deployment process provided by one embodiment of the present invention;

[0025] Figure 7 This is a structural diagram of a modeling and simulation device for on-orbit deployment of a ring reflector antenna provided by one embodiment of the present invention;

[0026] Figure 8 This is a hardware architecture diagram of a computer device provided by one embodiment of the present invention. DETAILED DESCRIPTION

[0027] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0028] As mentioned above, in related technologies, the disturbance of satellite attitude is usually analyzed based on the force, torque and stress data of the ring truss mesh parabolic antenna in the ground deployment test. However, this method has a large amount of calculation and a complex calculation process.

[0029] Based on this, the concept of the present invention is to equate the antenna deployment process to the motion process of multiple discrete rigid bodies based on the measured data of the geometric centroid change of the actual ring truss antenna during ground deployment, and calculate the attitude motion of the antenna deployment process by considering the attitude dynamics and kinematic equations of the multi-rigid body translational motion.

[0030] The specific implementation of the above concept is described below.

[0031] Please refer to Figure 1 An embodiment of the present invention provides a modeling and simulation method for on-orbit deployment of a ring reflector antenna, the method comprising:

[0032] Step 100, calculating the geometric displacement curves of the discretized equivalent rigid bodies of the ring truss antenna according to the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground;

[0033] Step 102: establishing a differential equation of motion of the reflector antenna for multi-equivalent rigid body translational motion based on the geometric displacement curve and the geometric installation relationship of the reflector antenna;

[0034] Step 104 : numerically integrate the motion differential equation according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the reflector antenna deployment process.

[0035] In an embodiment of the present invention, first, based on the geometric centroid change curve of the satellite reflector antenna during ground deployment, the geometric displacement curves of each discretized equivalent rigid body of the annular truss antenna are calculated; then, based on the geometric displacement curves of each discretized equivalent rigid body of the annular truss antenna and the geometric relationship of the annular truss antenna's installation on the entire satellite, the attitude dynamics and kinematic equations considering the translational motion of multiple rigid bodies are numerically integrated to obtain the satellite attitude motion changes during the reflector antenna deployment process. The calculation process used in this method is simple and effective, and has strong practicality. It can also be extended to the analysis of disturbances caused by other motion accessories on the satellite attitude, and is also of reference value for the simulation analysis of the in-orbit deployment process of other large satellites.

[0036] Described below Figure 1 How to perform the steps shown.

[0037] First, for step 100, the geometric displacement curves of the discretized equivalent rigid bodies of the annular truss antenna are calculated based on the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground.

[0038] In an embodiment of the present invention, the geometric displacement curve is determined by the following steps: fitting the geometric centroid displacement curve of the reflector antenna during ground deployment to obtain a sine polynomial function for geometric centroid displacement along the three axes of the coordinate system; and calculating, based on the sine polynomial function, the geometric displacement curve of each equivalent rigid body after the reflector antenna is discretized into multiple equivalent rigid bodies.

[0039] Specifically, according to Figure 2 The geometric configuration and coordinate system diagram of the large annular reflector antenna and satellite center body shown in the figure are as follows. Figure 3 The geometric centroid displacement curve of the reflector antenna shown in the figure is fitted during the ground deployment process, and the following is obtained: Figure 4 The sine polynomial function shown is calculated as:

[0040]

[0041] Where, is the geometric centroid displacement of the reflector antenna along the X, Y, and Z axes in the local coordinate system of the antenna during ground deployment at time t, and the unit is m.

[0042] Furthermore, the reflector antenna is discretized into an equivalent n according to the geometric centroid displacement calculation. AT After the rigid bodies are formed, the geometric displacement curves of each rigid body are:

[0043]

[0044] Where n AT is the number of equivalent rigid bodies discretized into the reflector antenna; is the geometric displacement of the i-th equivalent rigid body discretized at time t in the local coordinate system of the antenna, in units of m, i = 1,...,n AT

[0045] Then, for step 102, a motion differential equation of the multi-equivalent rigid body translational motion of the reflector antenna is established according to the geometric displacement curve and the geometric installation relationship of the reflector antenna.

[0046] In the embodiment of the present invention, the differential equation of motion is determined by the following steps:

[0047] Calculating the nominal displacement vector of each equivalent rigid body relative to the center of mass of the satellite central body in the satellite body coordinate system according to the geometric displacement curve of each equivalent rigid body;

[0048] The system dynamics matrix of the satellite is established based on the satellite's mass, the moment of inertia of the satellite's central body relative to the satellite's central body's center of mass, and the attitude transfer matrix of the satellite's system relative to the inertial system.

[0049] Establishing a system state vector matrix of the entire satellite according to the nominal displacement vector, and establishing a nonlinear term vector matrix of the entire satellite according to the velocity vector derived from the nominal displacement vector;

[0050] Establishing a control input vector matrix for the entire satellite according to the thrust acting on the reflector antenna equivalent rigid body in the satellite body coordinate system;

[0051] The motion differential equation is established according to the system dynamics matrix A, the system state vector matrix X, the control input vector matrix U and the nonlinear term vector matrix F:

[0052]

[0053] In the formula, A is the system dynamics matrix; X is the system state vector matrix; U is the control input vector matrix; F is the nonlinear term vector matrix. Specifically, according to the discretized n at time t AT The geometric displacement of an equivalent rigid body in the local coordinate system of the antenna is calculated. AT The nominal displacement of an equivalent rigid body relative to the center of mass of the satellite body in the satellite body coordinate system is:

[0054]

[0055] in, C is the displacement vector of the origin of the local coordinate system of the antenna relative to the center of mass of the satellite body in the satellite body coordinate system, in meters; ATB is the transfer matrix of the antenna local coordinate system relative to the satellite body coordinate system; is the nominal displacement vector of the i-th equivalent rigid body of the ring truss antenna relative to the center of mass of the satellite central body in the satellite body coordinate system at time t, in meters.

[0056] Furthermore, the system dynamics matrix A is established according to the nominal displacement vector as shown in the following formula:

[0057]

[0058] Where, () × Compute the matrix for the vector cross product; To discretize the reflector antenna into equivalent rigid bodies 1, 2, ..., n AT The number of units, kg; m B is the mass of the entire satellite (including the satellite center body and the ring truss antenna), in kg. The specific calculation formula is:

[0059]

[0060] Where m Cis the mass of the satellite's central body, in kg.

[0061] J B The moment of inertia of the entire satellite relative to the center of mass of the satellite body, in kgm 2 , the specific calculation formula is:

[0062]

[0063] Where, J CB The moment of inertia of the satellite center body relative to the center of mass of the satellite center body in the satellite body coordinate system, in kgm 2 ; State variables is the displacement vector of the ith equivalent rigid body of the annular truss antenna relative to the center of mass of the satellite body in the satellite body coordinate system, in meters.

[0064] C BI is the attitude transfer matrix of the satellite system relative to the inertial system; the matrix S ATC , in kgm, can be expressed as:

[0065]

[0066] Furthermore, the system state vector matrix X is established as shown in the following formula:

[0067]

[0068] Where, is the displacement of the satellite's center of mass relative to the inertial system, in meters; Θ BI is the representation parameter of the satellite body coordinate system relative to the inertial system; state variable is the displacement vector of the ith equivalent rigid body of the annular truss antenna relative to the center of mass of the satellite body in the satellite body coordinate system, in meters.

[0069] First-order rate of change of state variables is (3n AT +6)×1-dimensional matrix, specifically expressed as

[0070]

[0071] Where, is the first-order rate of change of the satellite's mass center relative to the inertial system, in m / s; It is the angular velocity of the satellite body coordinate system relative to the inertial system expressed in the satellite body system, in rad / s; is the velocity vector of the ith equivalent rigid body of the annular truss antenna relative to the center of mass of the satellite body in the satellite body coordinate system, in m / s.

[0072] Second-order rate of change of state variables is (3n AT +6)×1-dimensional matrix, specifically expressed as:

[0073]

[0074] Where, is the second-order rate of change of the satellite's center of mass relative to the inertial system, in m / s 2 ; It is the angular acceleration of the satellite body coordinate system relative to the inertial system in the satellite body system, in rad / s 2 ; is the acceleration vector of the ith equivalent rigid body of the ring truss antenna relative to the center of mass of the satellite body in the satellite body coordinate system, in m / s 2 .

[0075] In the embodiment of the present invention, the following formula is used to establish (3n AT +6)×1-dimensional control input vector matrix U:

[0076]

[0077] Where, is the three-axis thrust acting on the satellite center body in the inertial coordinate system, unit is Nm; It is the three-axis moment acting on the satellite center body in the satellite body coordinate, unit is Nms.

[0078] is the thrust acting on the i-th equivalent rigid body of the ring truss antenna in the satellite body coordinate system, in Nm. The specific calculation formula is:

[0079]

[0080] Where, Ω AT ,ξ AT is the tracking response angular frequency and damping ratio of the equivalent rigid body motion.

[0081] In the embodiment of the present invention, the nonlinear term vector matrix F is established by the following formula:

[0082]

[0083] Where, F i is the i-th component of the nonlinear term vector matrix F.

[0084] The component F1 of the matrix F can be expressed as

[0085]

[0086] The component F2 of the matrix F can be expressed as

[0087]

[0088] Component F of matrix F i+2 (i=1,...,n AT ), which can be expressed as

[0089]

[0090] With respect to step 104 , the motion differential equation is numerically integrated according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the reflector antenna deployment process.

[0091] In this embodiment of the present invention, a fourth-order Runge-Kutta method is used to numerically integrate the differential equation of motion constructed in the above process. Based on the integral calculation results, the change in the satellite attitude angle during the on-orbit deployment of the ring reflector antenna is obtained, which is used to analyze the disturbance of the antenna deployment on the satellite attitude. The specific calculation process is well known to those skilled in the art and will not be detailed here.

[0092] The curves of satellite roll, pitch and yaw attitude angle changes during the antenna deployment process obtained by this method are as follows: Figure 5 As shown in the figure, the angular velocity change curves of the satellite along the X-axis, Y-axis and Z-axis of the satellite during the antenna deployment process are as follows: Figure 6 shown.

[0093] Please refer to Figure 7 An embodiment of the present invention provides a modeling and simulation device for on-orbit deployment of a ring reflector antenna, the device comprising:

[0094] The calculation module 700 is used to calculate the geometric displacement curves of the discretized equivalent rigid bodies of the ring truss antenna according to the geometric centroid change curve of the satellite reflector antenna during the ground deployment process;

[0095] A modeling module 702 is configured to establish a differential equation of motion of a multi-equivalent rigid body translational motion of the reflector antenna based on the geometric displacement curve and the geometric installation relationship of the reflector antenna;

[0096] The determination module 704 is configured to perform numerical integration on the motion differential equation according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the deployment of the reflector antenna.

[0097] In the embodiment of the present invention, the calculation of the geometric displacement curves of the discretized equivalent rigid bodies of the ring truss antenna according to the geometric centroid change curve of the reflector antenna during its deployment on the ground includes:

[0098] Fitting a geometric centroid displacement curve of the reflector antenna during ground deployment to obtain a sine polynomial function of geometric centroid displacement along three axes of a coordinate system;

[0099] After the reflector antenna is discretized into a plurality of equivalent rigid bodies, a geometric displacement curve of each equivalent rigid body is calculated according to the sine polynomial function.

[0100] In an embodiment of the present invention, establishing the differential equation of motion of the reflector antenna for multi-equivalent rigid body translational motion based on the geometric relationship of the annular truss antenna installation according to the geometric displacement curve includes:

[0101] Calculating the nominal displacement vector of each equivalent rigid body relative to the center of mass of the satellite central body in the satellite body coordinate system according to the geometric displacement curve of each equivalent rigid body;

[0102] The system dynamics matrix of the satellite is established based on the satellite's mass, the moment of inertia of the satellite's central body relative to the satellite's central body's center of mass, and the attitude transfer matrix of the satellite's system relative to the inertial system.

[0103] Establishing a system state vector matrix of the entire satellite according to the nominal displacement vector, and establishing a nonlinear term vector matrix of the entire satellite according to the velocity vector derived from the nominal displacement vector;

[0104] Establishing a control input vector matrix for the entire satellite according to the thrust acting on the reflector antenna equivalent rigid body in the satellite body coordinate system;

[0105] The motion differential equation is established according to the system dynamics matrix, the system state vector matrix, the control input vector matrix and the nonlinear term vector matrix:

[0106]

[0107] Where A is the system dynamics matrix; X is the system state vector matrix; U is the control input vector matrix; and F is the nonlinear term vector matrix.

[0108] In the embodiment of the present invention, the system dynamics matrix is ​​established by the following formula:

[0109]

[0110] in,() × Calculate the matrix for vector cross product; m i is the mass of the i-th equivalent rigid body, i = 1, 2, ..., n AT ;m B is the mass of the whole star; J B is the moment of inertia of the entire satellite relative to the center of mass of the satellite; CBI is the attitude transfer matrix of the satellite system relative to the inertial system; is the nominal displacement vector of the i-th equivalent rigid body of the reflector antenna relative to the center of mass of the satellite central body.

[0111] In the embodiment of the present invention, the system state vector matrix is ​​established by the following formula:

[0112]

[0113] Where, is the displacement of the satellite's center of mass relative to the inertial system; Θ BI is the representation parameter of the satellite body coordinate system relative to the inertial system; is the nominal displacement vector.

[0114] In the embodiment of the present invention, the control input vector matrix is ​​established by the following formula:

[0115]

[0116] Where, is the three-axis thrust acting on the satellite center body in the inertial coordinate system; is the three-axis moment acting on the satellite center body in the satellite body coordinate; is the thrust acting on the i-th equivalent rigid body of the ring truss antenna in the satellite body coordinate system.

[0117] In the embodiment of the present invention, the nonlinear term vector matrix is ​​established by the following formula:

[0118]

[0119] Where, F i is the i-th component of the nonlinear term vector matrix F.

[0120] It should be noted that the modeling and simulation device for the on-orbit deployment of a circular reflector antenna provided in the above embodiment is only illustrated by the division of the above-mentioned functional modules. In actual applications, the above-mentioned functions can be assigned to different functional modules as needed, that is, the internal structure of the device can be divided into different functional modules to complete all or part of the functions described above. In addition, the modeling and simulation device for the on-orbit deployment of a circular reflector antenna provided in the above embodiment and the modeling and simulation method embodiment for the on-orbit deployment of a circular reflector antenna are based on the same concept. The specific implementation process is detailed in the method embodiment and will not be repeated here.

[0121] The embodiment of the present application also provides a computer device, please refer to Figure 8The computer device includes a processor and a memory, wherein the memory stores at least one instruction, at least one program, code set or instruction set, and the at least one instruction, at least one program, code set or instruction set is loaded and executed by the processor to implement the modeling and simulation method for the on-orbit deployment of the ring reflector antenna provided by the above-mentioned method embodiments.

[0122] An embodiment of the present application also provides a computer-readable storage medium, which stores at least one instruction, at least one program, code set or instruction set, and the at least one instruction, at least one program, code set or instruction set is loaded and executed by a processor to implement the modeling and simulation method for the in-orbit deployment of the annular reflector antenna provided in the above-mentioned method embodiments.

[0123] An embodiment of the present application also provides a computer program product, which includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium, and the processor executes the computer program, so that the computer device executes the modeling and simulation method for in-orbit deployment of a ring reflector antenna described in any of the above embodiments.

[0124] For the convenience of description, the above systems or devices are described as being divided into various modules or units according to their functions. Of course, when implementing the present application, the functions of each unit can be implemented in the same or multiple software and / or hardware.

[0125] Through the description of the above embodiments, it can be seen that those skilled in the art can clearly understand that the present application can be implemented by means of software plus a necessary general hardware platform. Based on this understanding, the technical solution of the present application, or the part that contributes to the prior art, can be embodied in the form of a software product, which can be stored in a storage medium such as ROM / RAM, a magnetic disk, an optical disk, etc., and includes a number of instructions for enabling a computer device (which can be a personal computer, a server, or a network device, etc.) to execute the methods described in various embodiments of the present application or certain parts of the embodiments.

[0126] Finally, it should be noted that, in this document, relational terms such as first, second, third, and fourth are used only to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprises," "comprising," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus comprising a series of elements includes not only those elements, but also other elements not explicitly listed, or elements inherent to such process, method, article, or apparatus. In the absence of further limitations, an element defined by the phrase "comprising a ..." does not exclude the presence of additional identical elements in the process, method, article, or apparatus comprising the element.

[0127] The above is only a preferred embodiment of the present application. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present application. These improvements and modifications should also be regarded as the scope of protection of the present application.

Claims

1. A modeling and simulation method for on-orbit deployment of a ring reflector antenna, characterized in that: The method comprises: According to the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground, the geometric displacement curves of the discretized equivalent rigid bodies of the ring truss antenna are calculated. Establishing a differential equation of motion for the multi-equivalent rigid body translational motion of the reflector antenna based on the geometric displacement curve and the geometric installation relationship of the reflector antenna; The motion differential equation is numerically integrated according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the deployment of the reflector antenna.

2. The method according to claim 1, wherein The calculation of the geometric displacement curves of the discretized equivalent rigid bodies of the ring truss antenna according to the geometric centroid change curve of the reflector antenna during its deployment on the ground includes: Fitting a geometric centroid displacement curve of the reflector antenna during ground deployment to obtain a sine polynomial function of geometric centroid displacement along three axes of a coordinate system; After the reflector antenna is discretized into a plurality of equivalent rigid bodies, a geometric displacement curve of each equivalent rigid body is calculated according to the sine polynomial function.

3. The method according to claim 1, wherein The geometric relationship of the ring truss antenna installation according to the geometric displacement curve is used to establish the motion differential equation of the multi-equivalent rigid body translation motion of the reflector antenna, including: Calculating the nominal displacement vector of each equivalent rigid body relative to the center of mass of the satellite central body in the satellite body coordinate system according to the geometric displacement curve of each equivalent rigid body; The system dynamics matrix of the satellite is established based on the satellite's mass, the moment of inertia of the satellite's central body relative to the satellite's central body's center of mass, and the attitude transfer matrix of the satellite's system relative to the inertial system. Establishing a system state vector matrix of the entire satellite according to the nominal displacement vector, and establishing a nonlinear term vector matrix of the entire satellite according to the velocity vector derived from the nominal displacement vector; Establishing a control input vector matrix for the entire satellite according to the thrust acting on the reflector antenna equivalent rigid body in the satellite body coordinate system; The motion differential equation is established according to the system dynamics matrix, the system state vector matrix, the control input vector matrix and the nonlinear term vector matrix: Where A is the system dynamics matrix; X is the system state vector matrix; U is the control input vector matrix; and F is the nonlinear term vector matrix.

4. The method according to claim 3, wherein The system dynamics matrix is ​​established by the following formula: in,() × Calculate the matrix for vector cross product; m i is the mass of the i-th equivalent rigid body, i = 1, 2, ..., n AT ;m B is the mass of the whole star; J B is the moment of inertia of the entire satellite relative to the center of mass of the satellite; C BI is the attitude transfer matrix of the satellite system relative to the inertial system; is the nominal displacement vector of the i-th equivalent rigid body of the reflector antenna relative to the center of mass of the satellite central body.

5. The method according to claim 3, wherein The system state vector matrix is ​​established by the following formula: Where, is the displacement of the satellite's center of mass relative to the inertial system; Θ BI is the representation parameter of the satellite body coordinate system relative to the inertial system; is the nominal displacement vector.

6. The method according to claim 3, wherein The control input vector matrix is ​​established by the following formula: Where, is the three-axis thrust acting on the satellite center body in the inertial coordinate system; is the three-axis moment acting on the satellite center body in the satellite body coordinate; is the thrust acting on the i-th equivalent rigid body of the ring truss antenna in the satellite body coordinate system.

7. The method according to claim 3, wherein The nonlinear term vector matrix is ​​established by the following formula: Where, F i is the i-th component of the nonlinear term vector matrix F.

8. A modeling and simulation device for on-orbit deployment of a ring reflector antenna, characterized in that: The device comprises: A calculation module is used to calculate the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna according to the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground; A modeling module, configured to establish a motion differential equation of a multi-equivalent rigid body translational motion of the reflector antenna based on the geometric displacement curve and the geometric installation relationship of the reflector antenna; The determination module is used to perform numerical integration on the motion differential equation according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the change in satellite attitude motion during the deployment of the reflector antenna.

9. A computer device, characterized in that: The computer device includes a memory and a processor, the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to implement the steps of any one of the methods described in claims 1-6.

10. A computer-readable storage medium, characterized in that The storage medium stores a computer program, which, when executed by a processor, implements the steps of the method according to any one of claims 1 to 6.

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