A method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects
By establishing a satellite finite element model and integrating multiple field effects, the problem of difficulty in predicting satellite antenna patterns in traditional methods was solved, and high-precision on-orbit electrical performance analysis and simulation were achieved.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- CHINA ACADEMY OF SPACE TECHNOLOGY
- Filing Date
- 2022-11-16
- Publication Date
- 2026-05-26
AI Technical Summary
Traditional decoupling methods are difficult to adapt to high-throughput satellites and mobile communication satellites with large capacity and smaller spot beams. They also have low computational efficiency under multi-field coupling and cannot accurately predict satellite electrical performance.
A finite element model of the satellite structure was established, and the model order was reduced. Combining multiple field effects such as heat, force, and light, the virtual power principle and the generalized force method of multi-field coupling were used to integrate multiple field effects into the same model, and the on-orbit antenna pattern of the satellite was calculated.
It achieves high-precision prediction of satellite electrical performance under multi-field coupling environment, solves the difficulty of dynamic prediction of antenna pattern under attitude change and structural deformation, and provides an efficient simulation analysis method.
Smart Images
Figure CN115859709B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of satellite overall design technology, and in particular, it is a method for determining the on-orbit radiation pattern of a satellite antenna that takes into account multiple field effects. Background Technology
[0002] The antenna patterns of spacecraft such as high-throughput communication satellites, high-capacity mobile communication satellites, and high-resolution and very-high-resolution space-based SAR satellites have a significant impact on the overall satellite payload performance, directly determining key indicators such as communication capacity, resolution, and geometric positioning accuracy. Accurately predicting the antenna patterns of these satellites is of great value for verifying the overall satellite design specifications and iteratively improving the overall satellite design scheme.
[0003] High-throughput communication satellites, high-capacity mobile communication satellites, and high-resolution and very-high-resolution space-based SAR satellites often carry large-aperture, high-precision antennas and large-area flexible solar arrays. The on-orbit radiation patterns of these antennas are highly susceptible to various factors. Satellites with flexible antennas are subject to external forces such as Earth's gravity, thruster exhaust, light pressure, and electromagnetic radiation, as well as internal forces such as dynamic imbalances in reaction wheels or control moment gyroscopes, rotation of the solar array drive mechanism, transient thermal shocks from entering and exiting the Earth's shadow, and propellant sloshing in the tanks. These factors cause vibrations in the large antennas themselves, changes in the relative position of the antenna and the feed source, and lead to antenna line-of-sight deviation or jitter, antenna pattern distortion, resulting in decreased satellite imaging or communication quality, and even structural damage to the satellite, severely impacting its performance.
[0004] Traditional methods generally employ decoupling approaches to solve antenna pattern problems from the antenna perspective. However, these methods are ill-suited for high-capacity, small-beam, very high-throughput satellites and mobile communication satellites, as well as the design iteration requirements of very high-resolution radar satellites for user performance verification and antenna pattern prediction. Traditional methods typically utilize multiple commercial software programs based on decoupling to solve the multi-field problems of satellite forces, heat, and light. Essentially, this artificially decouples multi-field coupling into a single-field problem, making it difficult to meet the demands of solving multi-physics coupling problems for high-precision satellites with large, flexible structures. Summary of the Invention
[0005] The technical problem solved by this invention is to overcome the shortcomings of existing technologies and address the difficulty in dynamically predicting antenna patterns caused by the disconnect between orbit, attitude control, thermal deformation, and antenna pattern prediction in traditional decoupling methods. This invention provides a method for determining the on-orbit antenna pattern considering multiple field effects, which solves the problems of difficulty in predicting satellite electrical performance under the coupling factors of continuous attitude changes, structural deformation, and thermal analysis, and low computational efficiency caused by single-field coupling. This provides an analytical method and implementation means for accurately predicting the on-orbit electrical performance of satellites with large flexible payloads.
[0006] The technical solution of this invention is:
[0007] A method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects includes:
[0008] Establish a finite element model of the satellite structure;
[0009] The finite element model of the on-board flexible structure was reduced in order to obtain the reduced dynamic model of the satellite.
[0010] Establish an on-orbit thermal analysis model for the satellite;
[0011] An on-orbit thermal analysis model of the satellite was analyzed to obtain the temperature of each node in the thermal analysis model. The temperature of each node in the on-orbit thermal analysis model was then mapped to the reduced-order dynamic model of the satellite to obtain the temperature, thermal strain, and thermal stress of each node in the reduced-order dynamic model, forming a node temperature load matrix, which was applied as an internal force to the reduced-order structural dynamic model of the satellite. A solar radiation force model of the satellite was established to obtain the solar radiation force matrix acting on each component of the satellite. Through the principle of virtual power, the solar radiation force was applied as an external force to the nodes of the reduced-order dynamic model of the satellite.
[0012] The gravitational forces exerted on each object of the satellite by the Earth and the Moon are calculated and applied as the external gravitational load matrix to the center of mass of each component in the reduced-order dynamic model of the satellite.
[0013] The satellite's flexible dynamic equations are determined using the nodal temperature load matrix, the solar radiation force matrix, and the gravitational external load matrices of various objects on the satellite.
[0014] Numerical integration of the satellite's flexible dynamics equations yields the generalized coordinates of the satellite's dynamics equations at the next time step.
[0015] Based on the generalized coordinates of the satellite dynamic equations at the new time step, and the flexible structure mode shape corresponding to the satellite's reduced-order dynamic model, the positions of each point on the antenna finite element model at the next time step are determined.
[0016] Based on the positions of each point on the antenna finite element model at the next moment, the antenna profile is fitted, and the far-field radiation pattern of the antenna corresponding to the antenna profile is calculated.
[0017] Preferably, the on-board flexible structure includes a solar array and an antenna.
[0018] Preferably, the principle of inertial completeness is adopted to reduce the order of the finite element model of the on-board flexible structure.
[0019] Preferably, the antenna profile is fitted using Zernike polynomials.
[0020] Preferably, the satellite flexible dynamics equations are as follows:
[0021]
[0022] Where M is the generalized mass matrix of the satellite system; X represents the generalized coordinates of the satellite system, and F... GenForce F represents the perturbation matrix of the satellite thruster. therm F represents the nodal temperature load matrix. sol F represents the satellite's solar radiative force matrix. grav This represents the gravitational external load matrix.
[0023] Preferably, the satellite solar radiation force model is established based on the geometric parameters, absorptivity, transmittance, and reflectivity of the satellite body, solar array, and antenna.
[0024] Preferably, the gravitational forces exerted on each object of the satellite by the Earth and the Moon are determined based on the mass, inertia characteristics of each satellite component and the topology of the satellite system, and using the universal gravitation formula.
[0025] The advantages of this invention compared to the prior art are:
[0026] Compared with existing methods, the method proposed in this invention integrates multiple field effects such as force, heat, and light into the same model through the equivalent generalized force method. This enables high-precision prediction of satellite electrical performance under space force, heat, and light environments with large flexible attachments. It avoids the problem that existing decoupling techniques cannot solve the problem of predicting satellite electrical performance under structural deformation, thermal deformation, and optical-pressure coupling under attitude changes. This provides an effective simulation means for the overall design and performance index analysis of satellites such as very high throughput communication satellites and SAR satellites with large flexible payloads. Attached Figure Description
[0027] Figure 1 This is a flowchart of the method of the present invention. Detailed Implementation
[0028] The present invention will become clearer and more apparent from the following detailed description. The present invention provides a method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects, as follows: Figure 1 As shown, the specific steps are as follows:
[0029] Step 1: Read in Coordinated Universal Time (UTC), satellite orbital root numbers, initial attitude and initial attitude angular velocity relative to the orbital coordinate system, initial temperature, geometric parameters, thermal emissivity, and thermal conductivity of each object on the satellite; read in parameters such as transmittance and reflectivity of the satellite body, solar array, antenna, etc., and the connection relationships between each component; read in the satellite topology, mass and inertia matrix of each object, and the flexible finite element model of the flexible solar array and antenna structure; establish a fully flexible multibody dynamics model of the satellite based on the general equations of dynamics.
[0030] Step 1.1 Based on the finite element model of the large flexible structure (such as solar panels and antennas) read in, obtain the elastic modulus matrix, mass and damping matrix of the object, as well as the initial position of each node in the finite element model;
[0031] Step 1.2 adopts the principle of inertial completeness to reduce the model order of the large flexible structure on the satellite (such as solar panels and antennas) to obtain the dynamic model of the satellite after the order reduction, and obtain the mode shape of the flexible structure after the order reduction;
[0032] x i =Φ i η i (1)
[0033] In the formula, x i Φ i η i These represent the physical coordinate system position of the node, the mode shape, and the generalized coordinates of the reduced-order model, respectively.
[0034] Step 1.3: Using the general equations of dynamics, establish the flexible dynamic equations of the satellite with a tree-like topology that considers multi-physics field effects:
[0035]
[0036] In the formula, M is the generalized mass matrix of the satellite system; X represents the generalized coordinates of the system, and F... GenForce F represents the perturbation array of satellite thrusters. therm F represents the generalized force matrix (i.e., the nodal temperature load matrix) related to thermal deformation. sol F represents the generalized external force array related to solar radiation (i.e., the satellite solar radiation force matrix). agrv This represents the Earth's gravitational array (i.e., the external gravitational load matrix) for each component on the star.
[0037] Step 2: Based on the second law of thermodynamics, establish an on-orbit thermal analysis model for the satellite.
[0038]
[0039] In the formula, ρ, c, k, T, t, q r Q represents the average density, heat capacity, thermal conductivity, temperature, time, radiative heat flux, and heat source of the satellite's main body panels, solar panels, antennas, and other components, respectively; these are discretized to establish a thermal analysis model of the entire satellite.
[0040]
[0041] In the formula, C and K e K rQ, T, and t represent the heat capacity matrix, heat conduction matrix, heat radiation matrix, heat flux load matrix, nodal temperature vector, and time, respectively.
[0042] Step 3: Combining the Coordinated Universal Time (UTC), satellite orbital root data, initial attitude and initial attitude angular velocity relative to the orbital coordinate system, initial temperature, geometric parameters, thermal radiation coefficient, and thermal conductivity coefficient of each object on the satellite, the on-orbit thermal analysis model of the satellite is analyzed to obtain the temperature of each node in the thermal analysis model. Using the element and node position relationship between the satellite thermal analysis model and the structural finite element model, the temperature of each node in the reduced-order finite element model of the satellite is interpolated; see Liu Guoqing, Luo Wenbo, and Tong Yelong, "Method for On-Orbit Full-Cycle Thermal Deformation Analysis of Spacecraft".
[0043] Step 4: Combining the reduced-order dynamic model of the satellite with the corresponding relationship between the node temperature, thermal strain, and thermal stress of the structural unit, a node temperature load matrix is formed, which is applied as an internal force to the reduced-order dynamic model of the satellite; see Liu Guoqing, Luo Wenbo, and Tong Yelong, "Method for On-orbit Full-Cycle Thermal Deformation Analysis of Spacecraft".
[0044] Step 5: Combine the geometric parameters, absorptivity, transmittance, reflectivity, and other parameters of the satellite body, solar array, and antenna read in Step 1 to establish a satellite solar radiation force model, and apply it as an external force to the satellite's reduced-order finite element model using the principle of virtual power.
[0045] Step 5.1 Utilize the satellite geometric model to obtain the reduced-order satellite geometric model. The graphical reduction method can be found in Deiml, Suderland, Reiss et al., "Development and evaluation of thermal model reduction algorithms for spacecraft";
[0046] Step 5.2: Combining the geometric parameters, transmittance, and reflectance of the main body, solar array, and antenna components read in Step 1, the solar radiation pressure formula is used to calculate the solar radiation force matrix acting on each component of the celestial body. This matrix is then applied to the centroids of the geometric elements on the outer surfaces of each component in the reduced-order finite element model of the satellite. The vector of the solar radiation force at the center of each geometric element can be calculated using the following formula.
[0047]
[0048] Step 5.3 Based on the principle of virtual power, the solar radiation force vector matrix acting on the geometric center of the outer surface of the object is transformed into external forces applied to the nodes of the satellite's reduced-order finite element model corresponding to the satellite body, solar array, and antenna; the generalized external forces are calculated based on the principle of virtual power. For details, please refer to Hong Jiazhen's "Computational Multibody Dynamics" P361.
[0049] Step Six: Combining the mass, inertia characteristics, and system topology of each satellite component read in Step One, calculate the gravitational forces exerted on each object of the satellite by the Earth and the Moon according to the universal gravitation formula, and apply them as the gravitational external load matrix to the center of mass of each satellite component; for the universal gravitation calculation formula, please refer to Liu Lin's "Orbital Mechanics of Artificial Earth Satellites" pp. 26-45.
[0050] Step 7: Using the nodal temperature load matrix calculated in Step 4, the satellite solar radiation force matrix obtained in Step 5, and the gravitational external load matrices of various objects on the satellite obtained in Step 6, determine the satellite's flexible dynamic equations.
[0051] Step 8: Perform numerical integration on the dynamic equations from Step 7 to obtain the generalized coordinates of the satellite dynamic equations at new times, and write the new integration time and generalized coordinates into the computer.
[0052] Step 9: Based on the generalized coordinates of the satellite dynamic equations obtained in Step 8, extract the modal coordinates of the antenna object, and combine them with the flexible structure mode shape obtained in Step 1.2 (i.e., the flexible structure mode shape corresponding to the satellite after the finite element model is reduced in order) to calculate the position of each point on the antenna finite element model at the next moment.
[0053] Step 10: Based on the positions of each point on the antenna calculated in Step 9, fit the antenna profile using a Zernike polynomial, and then use the far-field radiation pattern calculation formula of the deformed antenna described by the polynomial to calculate the far-field radiation pattern of the antenna profile. For details of the far-field radiation pattern calculation formula described by the Zernike polynomial, please refer to Shi Jiachen's "Electrical Performance Analysis and Profile Reconstruction Research of Reflector Antennas Based on Zernike Polynomials" pp. 12-16.
[0054] Step 11: Determine if the new integration time is less than the termination time. If the new integration time is less than the termination time, update the geometry of each satellite component according to the new generalized coordinates of the system dynamics equations at the new time, and return to Step 3; if it is greater than the termination time, the calculation terminates.
[0055] While the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the invention. Any person skilled in the art can make possible variations and modifications to the technical solutions of the present invention using the disclosed methods and techniques without departing from the spirit and scope of the invention. Therefore, any simple modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the content of the technical solutions of the present invention, shall fall within the protection scope of the present invention. Where there is no conflict, the embodiments of this application and the technical features thereof can be combined with each other.
[0056] The contents not described in detail in this specification are common knowledge to those skilled in the art.
Claims
1. A method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects, characterized in that, include: Establish a finite element model of the satellite structure; The finite element model of the on-board flexible structure was reduced in order to obtain the reduced dynamic model of the satellite. Establish an on-orbit thermal analysis model for the satellite; The on-orbit thermal analysis model of the satellite is analyzed to obtain the temperature of each node in the thermal analysis model. The temperature of each node in the on-orbit thermal analysis model of the satellite is mapped to the reduced-order dynamic model of the satellite to obtain the temperature, thermal strain and thermal stress of each node in the reduced-order dynamic model of the satellite, forming a node temperature load matrix, which is applied as internal force to the reduced-order structural dynamic model of the satellite. A solar radiation force model for the satellite is established to obtain the solar radiation force matrix acting on each component of the satellite. Then, using the principle of virtual power, the solar radiation force is applied as an external force to the nodes of the satellite's reduced-order dynamic model. The gravitational forces exerted on each object of the satellite by the Earth and the Moon are calculated and applied as the external gravitational load matrix to the center of mass of each component in the reduced-order dynamic model of the satellite. The satellite's flexible dynamic equations are determined using the nodal temperature load matrix, the solar radiation force matrix, and the gravitational external load matrices of various objects on the satellite. Numerical integration of the satellite's flexible dynamics equations yields the generalized coordinates of the satellite's dynamics equations at the next time step. Based on the generalized coordinates of the satellite dynamic equations at the new time step, and the flexible structure mode shape corresponding to the satellite's reduced-order dynamic model, the positions of each point on the antenna finite element model at the next time step are determined. Based on the positions of each point on the antenna finite element model at the next moment, the antenna profile is fitted, and the far-field radiation pattern of the antenna corresponding to the antenna profile is calculated.
2. The method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects according to claim 1, characterized in that: The on-board flexible structure includes: solar panels and antennas.
3. The method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects according to claim 1, characterized in that: The principle of inertial completeness is adopted to reduce the order of the finite element model of the on-board flexible structure.
4. The method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects according to claim 1, characterized in that: The antenna profile was fitted using Zernike polynomials.
5. A method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects according to any one of claims 1 to 4, characterized in that, The satellite flexible dynamics equations are as follows: Where M is the generalized mass matrix of the satellite system; X represents the generalized coordinates of the satellite system, and F... GenForce F represents the perturbation matrix of the satellite thruster. therm F represents the nodal temperature load matrix. sol F represents the satellite's solar radiative force matrix. grav This represents the gravitational external load matrix.
6. The method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects according to claim 5, characterized in that: The satellite solar radiation force model is established based on the geometric parameters, absorptivity, transmittance, and reflectivity of the satellite body, solar array, and antenna.
7. The method for determining the on-orbit radiation pattern of a satellite antenna considering multiple field effects according to claim 5, characterized in that: The gravitational forces exerted on each object of the satellite by the Earth and the Moon are determined based on the mass, inertia characteristics of each satellite component and the topology of the satellite system, and using the universal gravitation formula.