A method and device for modeling and simulating deployment of a ring reflector antenna in orbit

Abstract

The application discloses a modeling simulation method and device for on-orbit unfolding of a ring-shaped reflector antenna, and belongs to the field of spacecraft attitude control. The method comprises the following steps: calculating geometric displacement curves of each discretized equivalent rigid body of a ring-shaped truss antenna according to a geometric centroid change curve of a satellite reflector antenna in a ground unfolding process; establishing a motion differential equation of translational motion of the multi-equivalent rigid body of the reflector antenna according to the geometric displacement curves and a geometric installation relationship of the reflector antenna; and performing numerical integration on the motion differential equation according to a fourth-order Runge-Kutta method to determine a simulation model for representing the satellite attitude motion change in the unfolding process of the reflector antenna. The application can greatly reduce the simulation complexity of the on-orbit unfolding process of a large satellite.

Description

Technical Field

[0001] This invention relates to the field of spacecraft attitude control technology, and in particular to a modeling and simulation method and apparatus for on-orbit deployment of a ring reflector antenna. Background Technology

[0002] Geosynchronous orbit synthetic aperture radar (SAR) satellites typically employ large truss-type parabolic mesh antennas for microwave imaging. Because the antenna aperture is much larger than the launch vehicle's fairing envelope, the antenna must be retracted during launch and deployed after entering geosynchronous orbit. During the on-orbit deployment of the reflector antenna, the drastic changes in the overall satellite configuration, center of mass, and moment of inertia exert a significant impact on the satellite's control system. To analyze the effects of this impact, simulation analysis is needed based on the reflector deployment motion obtained from ground tests, combined with satellite attitude dynamics and kinematic modeling.

[0003] In related technologies, satellite attitude disturbances are usually analyzed based on force, torque, and stress data from ground deployment tests of the ring truss mesh parabolic antenna. However, this method involves a huge and complex amount of computation, which leads to its high cost and certain processing difficulties in practical applications.

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

[0005] This invention provides a modeling and simulation method and apparatus for 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] On the one hand, a modeling and simulation method for on-orbit deployment of a ring reflector antenna is provided, the method comprising:

[0007] Based on the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground, calculate the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna.

[0008] Based on the geometric displacement curve and the geometric installation relationship of the reflector antenna, the motion differential equations of the multi-equivalent rigid body translational motion of the reflector antenna are established.

[0009] The motion differential equations are numerically integrated using the fourth-order Runge-Kutta method to determine a simulation model that characterizes the changes in satellite attitude motion during the deployment of the reflector antenna.

[0010] On the other hand, a modeling and simulation device for on-orbit deployment of a ring reflector antenna is provided, the device comprising:

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

[0012] The modeling module is used to establish the differential equations of motion for the translational motion of the multi-equivalent rigid body 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 numerically integrate the motion differential equations according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the satellite attitude motion changes during the deployment of the reflector antenna.

[0014] On the other hand, a computer device is provided, the computer device including a memory and a processor, the memory for storing computer programs, and the processor for executing the computer programs stored in the memory to implement the steps of the above-described 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 a computer program is stored therein, and when the computer program is executed by a processor, the steps of the above-described modeling and simulation method for on-orbit deployment of a ring reflector antenna are implemented.

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

[0017] The technical solution provided by this invention can bring at least the following beneficial effects: First, based on the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground, the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna are calculated; then, based on the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna and the installation geometry of the ring truss antenna within the entire satellite, numerical integration is performed on the attitude dynamics and kinematic equations considering the translational motion of multiple rigid bodies 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 to satellite attitude caused by other moving attachments, and has reference value for the simulation analysis of the on-orbit deployment process of other large satellites. Attached Figure Description

[0018] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0019] Figure 1 This is a flowchart of a modeling and simulation method for on-orbit deployment of a ring reflector antenna according to an embodiment of the present invention;

[0020] Figure 2 This is a schematic diagram of the geometric configuration and coordinate system of a large ring reflector antenna and satellite center body provided in an embodiment of the present invention;

[0021] Figure 3 This is a schematic diagram of the geometric centroid displacement curves along the X, Y, and Z axes in the antenna local coordinate system during the ground deployment of a reflector antenna according to an embodiment of the present invention.

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

[0023] Figure 5 This is a schematic diagram of the satellite roll, pitch, and yaw attitude angle changes during the antenna deployment process provided in an embodiment of the present invention;

[0024] Figure 6 This is a schematic diagram of the attitude angular velocity changes of a satellite along the X-axis, Y-axis, and Z-axis of the celestial body during the antenna deployment process, provided by an 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 in an embodiment of the present invention;

[0026] Figure 8 This is a hardware architecture diagram of a computer device provided in an embodiment of the present invention. Detailed Implementation

[0027] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

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

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

[0030] The following describes the specific implementation of the above concept.

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

[0032] Step 100: Calculate the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna based on the geometric centroid change curve during the deployment of the satellite reflector antenna on the ground.

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

[0034] Step 104: Perform numerical integration on the motion differential equations according to the fourth-order Runge-Kutta method to determine the simulation model used to characterize the changes in satellite attitude motion during the deployment of the reflector antenna.

[0035] In this embodiment of the invention, firstly, based on the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground, the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna are calculated. Then, based on the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna and the installation geometry of the ring truss antenna within the entire satellite, numerical integration is performed on the attitude dynamics and kinematic equations considering the translational motion of multiple rigid bodies 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 to satellite attitude caused by other moving attachments, and has reference value for the simulation analysis of the on-orbit deployment process of other large satellites.

[0036] The following description Figure 1 The execution method for each step is shown.

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

[0038] In this embodiment of the invention, the geometric displacement curve is determined by the following steps: the geometric centroid displacement curve of the reflector antenna during its deployment on the ground is fitted to obtain a sinusoidal polynomial function of geometric centroid displacement along the three axes of the coordinate system; the geometric displacement curve of each equivalent rigid body is calculated based on the sinusoidal polynomial function after the reflector antenna is discretized into multiple equivalent rigid bodies.

[0039] Specifically, according to such Figure 2 The diagram shown illustrates the geometric configuration and coordinate system of the large ring reflector antenna and the satellite's central body. First, as shown... Figure 3 The geometric centroid displacement curve of the reflector antenna during its ground deployment process, as shown, is fitted to obtain the following: Figure 4 The formula for calculating the sinusoidal polynomial function shown is as follows:

[0040]

[0041] In the formula, Let t represent the geometric centroid displacements along the X, Y, and Z axes in the antenna's local coordinate system during the antenna's deployment on the ground, all in meters.

[0042] Furthermore, the reflector antenna is discretized into an equivalent n-value based on the geometric centroid displacement. AT After the rigid body is assembled, the geometric displacement curves of each rigid body are as follows:

[0043]

[0044] In the formula, n AT The number of equivalent rigid bodies to discretize the reflector antenna; Let be the geometric displacement of the i-th equivalent rigid body discretized at time t in the antenna local coordinate system, in meters (m), where i = 1, ..., n. AT

[0045] Then, for step 102, based on the geometric displacement curve and the geometric installation relationship of the reflector antenna, the motion differential equations of the multi-equivalent rigid body translational motion of the reflector antenna are established.

[0046] In this embodiment of the invention, the equation of motion is determined through the following steps:

[0047] Based on the geometric displacement curve of each equivalent rigid body, calculate the nominal displacement vector of the equivalent rigid body relative to the center of mass of the satellite in the satellite body coordinate system.

[0048] Based on the satellite's total mass, the moment of inertia of the satellite's central body relative to its center of mass, and the attitude transfer matrix of the satellite's own system relative to the inertial frame, the overall system dynamics matrix of the satellite is established.

[0049] Based on the nominal displacement vector, establish the overall system state vector matrix of the satellite, and establish the overall nonlinear term vector matrix of the satellite based on the velocity vector obtained by differentiating the nominal displacement vector.

[0050] Based on the thrust acting on the equivalent rigid body of the reflector antenna in the satellite body coordinate system, establish the overall control input vector matrix of the satellite.

[0051] The motion differential equation is established based on 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; and F is the nonlinear term vector matrix. Specifically, based on the discretized n at time t... AT Calculate the geometric displacement of an equivalent rigid body in the antenna's local coordinate system, n. AT The nominal displacement of an equivalent rigid body relative to the satellite's center of mass in the satellite's body coordinate system:

[0054]

[0055] in, C is the displacement vector of the origin of the antenna's local coordinate system relative to the satellite's center of mass in the satellite's body coordinate system, expressed in meters (m). ATB This is the transition matrix between the antenna local coordinate system and the satellite body coordinate system; Let be 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 in the satellite body coordinate system at time t, in meters.

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

[0057]

[0058] In the formula, () × Calculate the matrix for the vector cross product; To discretize the reflector antenna into equivalent rigid bodies 1, 2, ..., n AT The number of items, in kg; m B The mass of the entire satellite (including the satellite's central body and the ring truss antenna) is expressed in kg. The specific calculation formula is as follows:

[0059]

[0060] In the formula, m CThe mass of the satellite's central body is expressed in kg.

[0061] J B The moment of inertia of the entire satellite relative to its center of mass is expressed in kgm. 2 The specific calculation formula is as follows:

[0062]

[0063] In the formula, J CB The moment of inertia of the satellite's central body relative to its center of mass in the satellite's body coordinate system, expressed in kgm. 2 State variables Let be the displacement vector of the i-th equivalent rigid body of the ring truss antenna relative to the center of mass of the satellite in the satellite's body coordinate system, in meters.

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

[0065]

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

[0067]

[0068] In the formula, Θ represents the displacement of the satellite's center of mass relative to the inertial frame, in meters (m). BI The parameters representing the satellite's body coordinate system relative to the inertial frame; state variables. Let be the displacement vector of the i-th equivalent rigid body of the ring truss antenna relative to the center of mass of the satellite in the satellite's body coordinate system, in meters.

[0069] First-order rate of change of state variable For (3n) AT A 6×1 dimensional matrix, specifically represented as

[0070]

[0071] In the formula, The first-order rate of change of the displacement of the satellite's center of mass relative to the inertial frame is expressed in m / s. This represents the angular velocity of the satellite's body coordinate system relative to the inertial frame in the satellite's own system, in rad / s; Let be the velocity vector of the i-th equivalent rigid body of the ring truss antenna relative to the center of mass of the satellite in the satellite's body coordinate system, in m / s.

[0072] Second rate of change of state variable For (3n) AT A 6×1 dimensional matrix, specifically represented as:

[0073]

[0074] In the formula, The second rate of change of displacement of the satellite's center of mass relative to the inertial frame, in m / s. 2 ; This represents the angular acceleration of the satellite's body coordinate system relative to the inertial frame in the satellite's own system, in rad / s. 2 ; Let be the acceleration vector of the i-th equivalent rigid body of the ring truss antenna relative to the center of mass of the satellite in the satellite's body coordinate system, in m / s². 2 .

[0075] In this embodiment of the invention, (3n) is established by the following formula. AT +6)×1 dimensional control input vector matrix U:

[0076]

[0077] In the formula, The thrust is the three-axis thrust acting on the satellite's central body in the inertial coordinate system, in Nm. The torque is the three-axis torque acting on the central body of the satellite in the satellite body coordinate system, in N / ms.

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

[0079]

[0080] In the formula, Ω AT ξ AT The tracking response angular frequency and damping ratio of the equivalent rigid body motion are given.

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

[0082]

[0083] In the formula, F i Let be the i-th component of the nonlinear term vector matrix F.

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

[0085]

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

[0087]

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

[0089]

[0090] For step 104, the motion differential equation is numerically integrated according to the fourth-order Runge-Kutta method to determine the simulation model used to characterize the satellite attitude motion changes during the deployment of the reflector antenna.

[0091] In this embodiment of the invention, the fourth-order Runge-Kutta method is used to numerically integrate the motion differential equations constructed in the above process. Based on the integration results, the changes in the satellite's attitude angles during the on-orbit deployment of the ring reflector antenna can be obtained, which is used to analyze the disturbance of the antenna deployment on the satellite's attitude. The specific calculation process is well known to those skilled in the art and will not be described in detail here.

[0092] The curves of satellite roll, pitch, and yaw attitude angle changes during antenna deployment obtained using this method are shown below. Figure 5 As shown, the satellite's attitude angular velocity changes along the X, Y, and Z axes of the celestial body during the antenna deployment process are as follows: Figure 6 As shown.

[0093] Please refer to Figure 7 This 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 each discretized equivalent rigid body of the ring truss antenna based on the geometric centroid change curve during the deployment of the satellite reflector antenna on the ground.

[0095] Modeling module 702 is used to establish the motion differential equations of 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.

[0096] The determination module 704 is used to perform numerical integration on the motion differential equation according to the fourth-order Runge-Kutta method to determine the simulation model used to characterize the satellite attitude motion changes during the deployment of the reflector antenna.

[0097] In this embodiment of the invention, calculating the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna based on the geometric centroid change curve during the deployment of the reflector antenna on the ground includes:

[0098] The geometric centroid displacement curve of the reflector antenna during its deployment on the ground is fitted to obtain a sinusoidal polynomial function of geometric centroid displacement along the three axes of the coordinate system.

[0099] The geometric displacement curve of each equivalent rigid body is calculated based on the sinusoidal polynomial function after the reflector antenna is discretized into multiple equivalent rigid bodies.

[0100] In this embodiment of the invention, establishing the differential equations of motion for the translational motion of the reflector antenna based on the geometric relationship of the annular truss antenna installation according to the geometric displacement curve includes:

[0101] Based on the geometric displacement curve of each equivalent rigid body, calculate the nominal displacement vector of the equivalent rigid body relative to the center of mass of the satellite in the satellite body coordinate system.

[0102] Based on the satellite's total mass, the moment of inertia of the satellite's central body relative to its center of mass, and the attitude transfer matrix of the satellite's own system relative to the inertial frame, the overall system dynamics matrix of the satellite is established.

[0103] Based on the nominal displacement vector, establish the overall system state vector matrix of the satellite, and establish the overall nonlinear term vector matrix of the satellite based on the velocity vector obtained by differentiating the nominal displacement vector.

[0104] Based on the thrust acting on the equivalent rigid body of the reflector antenna in the satellite body coordinate system, establish the overall control input vector matrix of the satellite.

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

[0106]

[0107] In the formula, 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 this embodiment of the invention, the system dynamics matrix is ​​established using the following formula:

[0109]

[0110] in,() × The matrix for calculating the cross product of vectors; m i Let be the mass of the i-th equivalent rigid body, i = 1, 2, ..., n AT m B For the whole satellite mass; J B C is the moment of inertia of the entire satellite relative to its center of mass;BI This represents the attitude transfer matrix of the satellite system relative to the inertial frame. Let be the nominal displacement vector of the i-th equivalent rigid body of the reflector antenna relative to the center of mass of the satellite.

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

[0112]

[0113] In the formula, Θ represents the displacement of the satellite's center of mass relative to the inertial frame; BI These are the parameters representing the satellite's body coordinate system relative to the inertial frame; Let be the nominal displacement vector.

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

[0115]

[0116] In the formula, The three-axis thrust acting on the satellite's central body in the inertial coordinate system; The three-axis torques acting on the satellite's central body in the satellite's body coordinates; It represents the thrust acting on the i-th equivalent rigid body of the ring truss antenna in the satellite body coordinate system.

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

[0118]

[0119] In the formula, F i Let be the i-th component of the nonlinear term vector matrix F.

[0120] It should be noted that the modeling and simulation device for on-orbit deployment of a ring reflector antenna provided in the above embodiments is only an example illustrating the division of the functional modules. In practical applications, the 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. Furthermore, the modeling and simulation device for on-orbit deployment of a ring reflector antenna provided in the above embodiments and the modeling and simulation method embodiments for on-orbit deployment of a ring reflector antenna belong to the same concept; the specific implementation process is detailed in the method embodiments and will not be repeated here.

[0121] Embodiments of this application also provide a computer device, please refer to... Figure 8The computer device includes a processor and a memory, the memory storing at least one instruction, at least one program, code set or instruction set, the at least one instruction, at least one program, code set or instruction set being loaded and executed by the processor to implement the modeling and simulation method for on-orbit deployment of the loop reflector antenna provided in the above-described method embodiments.

[0122] Embodiments of this application also provide a computer-readable storage medium storing at least one instruction, at least one program, code set, or instruction set, wherein 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 on-orbit deployment of a loop reflector antenna provided in the above-described method embodiments.

[0123] Embodiments of this application also provide 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 executes the computer program, causing the computer device to perform the modeling and simulation method for on-orbit deployment of a ring reflector antenna as described in any of the above embodiments.

[0124] For ease of description, the above systems or devices are described separately as various modules or units based on their functions. Of course, in implementing this application, the functions of each unit can be implemented in one or more software and / or hardware components.

[0125] As can be seen from the above description of the embodiments, those skilled in the art can clearly understand that this application can be implemented by means of software plus necessary general-purpose hardware platforms. Based on this understanding, the technical solution of this application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product can be stored in a storage medium, such as ROM / RAM, magnetic disk, optical disk, etc., and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute the methods described in various embodiments or some parts of the embodiments of this application.

[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, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0127] The above description is only a preferred embodiment of this application. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of this application, and these improvements and modifications should also be considered within the scope of protection of this application.

Claims

1. A modeling and simulation method for on-orbit deployment of a ring reflector antenna, characterized in that, The method includes: Based on the geometric centroid change curve of the satellite reflector antenna during its deployment on the ground, calculate the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna, including: The geometric centroid displacement curve of the reflector antenna during its deployment on the ground is fitted to obtain a sinusoidal polynomial function of geometric centroid displacement along the three axes of the coordinate system. The geometric displacement curve of each equivalent rigid body is calculated based on the sinusoidal polynomial function after the reflector antenna is discretized into multiple equivalent rigid bodies. Based on the geometric displacement curve and the geometric installation relationship of the reflector antenna, the differential equations of motion for the multi-equivalent rigid body translational motion of the reflector antenna are established, including: Based on the geometric displacement curve of each equivalent rigid body, calculate the nominal displacement vector of the equivalent rigid body relative to the center of mass of the satellite in the satellite body coordinate system. Based on the satellite's total mass, the moment of inertia of the satellite's central body relative to its center of mass, and the attitude transfer matrix of the satellite's own system relative to the inertial frame, the overall system dynamics matrix of the satellite is established. Based on the nominal displacement vector, establish the overall system state vector matrix of the satellite, and establish the overall nonlinear term vector matrix of the satellite based on the velocity vector obtained by differentiating the nominal displacement vector; Based on the thrust acting on the equivalent rigid body of the reflector antenna in the satellite body coordinate system, establish the overall control input vector matrix of the satellite. The motion differential equation is established based on the system dynamics matrix, the system state vector matrix, the control input vector matrix, and the nonlinear term vector matrix: In the formula, The system dynamics matrix; The system state vector matrix; To control the input vector matrix; It is a vector matrix of nonlinear terms; The motion differential equations are numerically integrated using the fourth-order Runge-Kutta method to determine a simulation model that characterizes the changes in satellite attitude motion during the deployment of the reflector antenna.

2. The method as described in claim 1, characterized in that, The system dynamics matrix is ​​established using the following formula: in, Calculate the matrix for the vector cross product; m i Let be the mass of the i-th equivalent rigid body, i = 1, 2, ... ; For the whole satellite mass; Let be the moment of inertia of the entire satellite relative to the center of mass of the satellite's central body; This represents the attitude transfer matrix of the satellite system relative to the inertial frame. For the reflector antenna number The nominal displacement vector of an equivalent rigid body relative to the center of mass of the satellite.

3. The method as described in claim 1, characterized in that, The system state vector matrix is ​​established using the following formula: In the formula, This represents the displacement of the satellite's center of mass relative to the inertial frame. These are the parameters representing the satellite's body coordinate system relative to the inertial frame; Let be the nominal displacement vector.

4. The method as described in claim 1, characterized in that, The control input vector matrix is ​​established using the following formula: In the formula, The three-axis thrust acting on the satellite's central body in the inertial coordinate system; The three-axis torques acting on the satellite's central body in the satellite's body coordinates; ( ) represents the thrust acting on the i-th equivalent rigid body of the ring truss antenna in the satellite body coordinate system.

5. The method as described in claim 1, characterized in that, The nonlinear term vector matrix is ​​established by the following formula: In the formula, Let be the i-th component of the nonlinear term vector matrix F.

6. A modeling and simulation device for on-orbit deployment of a ring reflector antenna, characterized in that, The apparatus, used in the method of any one of claims 1-5, comprises: The calculation module is used to calculate the geometric displacement curves of each discretized equivalent rigid body of the ring truss antenna based on the geometric centroid change curve during the deployment of the satellite reflector antenna on the ground. The modeling module is used to establish the differential equations of motion for the translational motion of the multi-equivalent rigid body 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 numerically integrate the motion differential equations according to the fourth-order Runge-Kutta method to determine a simulation model for characterizing the satellite attitude motion changes during the deployment of the reflector antenna.

7. A computer device, characterized in that, The computer device includes a memory and a processor. The memory is used to store computer programs, and the processor is used to execute the computer programs stored in the memory to implement the steps of the method according to any one of claims 1-5.

8. 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 described in any one of claims 1-5.

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