Decoupling mechanical test method for batch production of satellite solar wings

Abstract

The invention provides a decoupling mechanical test method for batch-produced satellite solar wings, which comprises the following steps: establishing a satellite finite element model, comparing responses of two states with and without solar wings, and then independently carrying out a mechanical test on satellites; performing a mechanical test on the solar wing independently; carrying out a mechanical test on the solar wing mounted satellite; evaluating the response of the solar wing with the solar wing and the response of the satellite in the non-solar wing state and the response of the satellite; evaluating a decoupling result; and according to the evaluation result, implementing the decoupling mechanical test of the batch-produced satellite solar wing, or terminating the implementation of the decoupling mechanical test of the batch-produced satellite solar wing. The method not only can meet the vibration test assessment of the satellite body and the solar wing, but also can reduce the solar wing installation frequency of the batch-produced satellites and improve the development efficiency of the batch-produced satellites.

Description

Technical Field

[0001] The present invention relates to a satellite mechanical test method, and specifically, to a decoupled mechanical test method for the solar wings of mass-produced satellites. Background Art

[0002] It is difficult to obtain the technical status of satellite tests. However, from the publicly released satellite mechanical test diagrams, it can be seen that satellites are all tested mechanically with solar wings. Currently, in satellite development, mechanical tests are all carried out with solar wings. For example, in the foreign standard "Test requirments for launch, upper-stage and space vehicles" (SMC-S-016), it is stipulated that before the start of system-level tests, all flight products and software should be installed. In the national military standard "Test Requirements for Launch Vehicles, Upper Stages and Spacecraft" (GJB 1027A-2020), the system-level test items, purposes and related test requirements are stipulated, and the test description of the vibration test describes the state of the aircraft as the launch state.

[0003] Mass-produced satellites have characteristics such as high overall state similarity, short development cycle, and fast launch requirement speed, which pose extremely high requirements for the satellite development process and development efficiency. However, in the traditional satellite development process, there are many processes for installing and electrically detonating the solar wings before and after the single-satellite mechanical test, involving many positions, long time periods, and high consumption of human and material resources. Summary of the Invention

[0004] Aiming at the deficiencies in the prior art and for the launch tasks of mass-produced satellites, the present invention proposes a decoupled mechanical test method for the solar wings of mass-produced satellites, which comprehensively designs the comprehensiveness of the process and the coverage and reliability of the test process, and can optimize the installation and disassembly times of the solar wings on the basis of effectively meeting the assessment and screening of satellite mechanical characteristics, improve the utilization efficiency of personnel and solar wing deployment equipment resources, and meet the requirements of mass-produced satellites.

[0005] The present application provides a decoupled mechanical test method for the solar wings of mass-produced satellites, including:

[0006] S1: Establish a first satellite finite element model and a second satellite finite element model, where the first satellite finite element model is a finite element model of the entire satellite with solar wings, and the second satellite finite element model is a finite element model of the satellite body;

[0007] S2: Conduct modal analysis and sine response analysis on the first satellite finite element model and the second satellite finite element model respectively to compare the differences in the stiffness simulation values of the satellite body and the differences in the acceleration response simulation values at the concerned positions in the two states with and without solar wings;

[0008] S3: Conduct three-directional sinusoidal characteristic-level swept-frequency tests on the satellite body to obtain the acceleration responses at the positions of interest on the satellite body.

[0009] S4: Conduct three-directional sinusoidal characteristic-level swept-frequency tests on the solar panel unit to obtain the acceleration responses at the positions of interest on the solar panel unit.

[0010] S5: Install the satellite solar panel on the satellite.

[0011] S6: Conduct three-directional sinusoidal characteristic-level swept-frequency tests on the satellite body equipped with the solar panel to obtain the acceleration responses at the same positions of interest on the satellite body as in step S3, and the acceleration responses at the same positions of interest on the solar panel unit as in step S4.

[0012] S7: Compare the acceleration responses at the positions of interest on the satellite body obtained in step S6 with those obtained in step S3, so as to obtain the differences in the acceleration responses at the positions of interest on the satellite body in the two states of with and without the solar panel.

[0013] S8: Compare the acceleration responses at the positions of interest on the solar panel unit obtained in step S6 with those obtained in step S4, so as to obtain the differences in the acceleration responses at the positions of interest on the solar panel unit in the two states of with and without the solar panel.

[0014] S9: Evaluate the decoupling consistency.

[0015] According to the method provided by an embodiment of the present application, in step S1, the first satellite finite element model is a finite element model established for the satellite body equipped with the solar panel, and the second satellite finite element model is a finite element model established for the satellite body without the solar panel.

[0016] According to the method provided by an embodiment of the present application, step S2 includes, in the finite element software, by conducting sinusoidal response analysis, calculating the simulated values of the acceleration responses at the positions of interest on the satellite body under the acceptance-level sinusoidal vibration excitation in the two states of with and without the solar panel, and obtaining the differences in the simulated values of the acceleration responses of the satellite body in the two states of with and without the solar panel.

[0017] According to the method provided by an embodiment of the present application, step S2 includes, in the finite element software, by conducting modal analysis, obtaining the simulated values of the stiffness of the satellite body in the two states of with and without the solar panel, and obtaining the differences in the simulated values of the stiffness of the satellite body in the two states of with and without the solar panel.

[0018] The method provided according to an embodiment of the present application, wherein in step S4, the measurement points of the solar wing single machine sensor are not less than those of the solar wing in the full satellite mechanical test, and cover the measurement points of the solar wing in the full satellite mechanical test.

[0019] The method provided according to an embodiment of the present application, wherein the acceleration response includes the peak value and the occurrence frequency of the peak value.

[0020] The method provided according to an embodiment of the present application, in step S9, if the decoupling consistency evaluation result is poor, it is determined that the decoupling mechanical test of the mass-produced satellite and the solar wing cannot be carried out subsequently.

[0021] The method provided according to an embodiment of the present application, in step S9, if the decoupling consistency evaluation result is relatively poor, it is determined that the decoupling mechanical test of the mass-produced satellite and the solar wing cannot be carried out subsequently; if the decoupling consistency evaluation result is relatively good, it is determined that the decoupling mechanical test of the mass-produced satellite and the solar wing can be carried out subsequently.

[0022] The method provided according to an embodiment of the present application, the evaluation indexes for evaluating the decoupling consistency in step S9 include:

[0023] The difference in the stiffness simulation values of the satellite body and the difference in the acceleration response simulation values at the concerned positions in the two states of having and not having the solar wing obtained in step S2;

[0024] The difference in the acceleration response at the concerned positions of the satellite body in the two states of having and not having the solar wing obtained in step S7;

[0025] The difference in the acceleration response at the concerned positions of the solar wing single machine in the two states of having and not having the solar wing obtained in step S8.

[0026] The present application also provides a computer-readable storage medium storing software instructions, and the software instructions, when executed, implement the above method.

[0027] This method can not only meet the vibration test requirements of the satellite body and the solar wing, but also reduce the number of times of installing the solar wing on the mass-produced satellite, and improve the development efficiency of the mass-produced satellite.

[0028] The beneficial effects of the present invention compared with the prior art are:

[0029] ① The present invention provides a mechanical test method for mass-produced satellites without solar wings.

[0030] ② The test method provided by the present invention optimizes the number of times of installing and disassembling the solar wing, improves the utilization efficiency of personnel and solar wing deployment equipment resources, and reduces the labor and material costs.

[0031] ③The test method of the present invention shortens the R & D cycle of mass-produced satellites, improves the launch speed of mass-produced satellites, and has broad application and popularization value. Brief Description of the Drawings

[0032] The above characteristics, technical features, advantages and their implementation manners of the present application will be further described below in a clear and understandable manner through the description of preferred embodiments and in combination with the drawings. The following drawings are only intended to illustrate and explain the present application schematically and do not limit the scope of the present application. Among them:

[0033] Figure 1 It is a flowchart of the decoupling mechanical test method for the solar wings of mass-produced satellites of the present invention.

[0034] Figure 2 It shows the development process of the decoupling mechanical test of mass-produced satellites and solar wings in the prior art. Detailed Embodiments

[0035] In order to have a clearer understanding of the technical features, purposes and effects of the present application, the detailed embodiments of the present application will now be described with reference to the drawings.

[0036] The present application provides a satellite mechanical test method for mass-produced satellites without solar wings, which optimizes the test processes before and after, such as Figure 2 As shown, compared with the traditional satellite development process, the installation, manual deployment and illumination test of the satellite and solar wings before the original mechanical test, the installation, testing and pressing of the pyrotechnic device, and the electro-explosive deployment, illumination test and removal of the solar wings after the mechanical test are optimized. When modifying in the launch state, the above work items of the solar wings are all tested, and the actual power supply capacity of the pyrotechnic device is carried out after the mechanical test and before this stage.

[0037] An embodiment of the present application provides a decoupling mechanical test method for the solar wings of mass-produced satellites, including the following steps:

[0038] S1: Establish a first satellite finite element model and a second satellite finite element model, where the first satellite finite element model is a finite element model of the whole satellite (i.e., a finite element model established for the satellite body equipped with solar wings), and the second satellite finite element model is a finite element model of the satellite body (i.e., a finite element model established for the satellite body without solar wings);

[0039] S2: Carry out modal analysis and sine response analysis on the first satellite finite element model and the second satellite finite element model respectively to compare the differences in the stiffness simulation values of the satellite body and the differences in the acceleration response simulation values at the concerned positions in the two states with and without solar wings.

[0040] Specifically, in the finite element software, through modal analysis, the stiffness simulation values of the satellite body in the two states with and without solar arrays are obtained, and the differences in the stiffness simulation values of the satellite body in the two states with and without solar arrays are obtained.

[0041] Specifically, in the finite element software, through sine response analysis, the acceleration response (including peak value and occurrence frequency) simulation values of the concerned positions of the satellite body under the acceptance-level sine vibration excitation in the two states with and without solar arrays are calculated, and the differences in the acceleration response simulation values of the satellite body in the two states with and without solar arrays are obtained. If the peak value and occurrence frequency of the satellite body response are approximately the same, it indicates that the installation of the solar array has little impact on the mechanical transfer characteristics of the satellite body and the risk is controllable.

[0042] S3: Conduct three-direction sine characteristic-level sweep frequency tests on the satellite body (excluding the solar array) to obtain the acceleration response of the concerned positions of the satellite body. The specific test method can be carried out with reference to QJ 1579A-2005;

[0043] S4: Conduct three-direction sine characteristic-level sweep frequency tests on the solar array unit. The number of sensor measuring points on the solar array unit is not less than that of the solar array in the full satellite mechanical test, and it covers the measuring points of the solar array in the full satellite mechanical test, to obtain the acceleration response of the concerned positions on the solar array unit;

[0044] S5: Install the satellite solar array on the satellite, and at the same time confirm that the installation of the sensors on the satellite and the solar array is correct;

[0045] S6: After the satellite installs the solar array, conduct three-direction sine characteristic-level sweep frequency tests on the full satellite (i.e., the satellite body equipped with the solar array) to obtain the acceleration response of the concerned positions of the satellite body same as in step S3, and the acceleration response of the concerned positions of the solar array unit same as in step S4;

[0046] S7: Compare the acceleration response of the concerned positions of the satellite body obtained in step S6 with the acceleration response of the concerned positions of the satellite body obtained in step S3, so as to obtain the difference in the acceleration response of the concerned positions of the satellite body in the two states with and without solar arrays;

[0047] S8: Compare the acceleration response of the concerned positions of the solar array unit obtained in step S6 with the acceleration response of the concerned positions of the solar array unit obtained in step S4, so as to obtain the difference in the acceleration response of the concerned positions of the solar array unit in the two states with and without solar arrays;

[0048] S9: Evaluate the decoupling consistency. If the evaluation result of the decoupling consistency is good, it is determined that the decoupling mechanical test of the production satellite and the solar wing can be carried out subsequently, and the specific test method can be referred to QJ 1579A-2005; if the evaluation result of the decoupling consistency is poor, it is determined that the decoupling mechanical test of the production satellite and the solar wing cannot be carried out subsequently. The evaluation indicators include:

[0049] The difference in the stiffness simulation values of the satellite body and the difference in the acceleration response simulation values at the concerned positions in the two states of the satellite body with and without the solar wing obtained in step S2;

[0050] The difference in the acceleration response at the concerned positions of the satellite body in the two states of the satellite body with and without the solar wing obtained in step S7;

[0051] The difference in the acceleration response at the concerned positions of the single solar wing in the two states of the single solar wing with and without the satellite body obtained in step S8.

[0052] Through the above evaluation indicators, it is possible to judge whether the mechanical test response of the single solar wing envelopes the magnitude of the whole satellite test, and it is also possible to judge whether the acceleration response at the concerned positions of the whole satellite without the solar wing can cover the whole satellite test state, and it can meet the launch vehicle and single machine test conditions.

[0053] The decoupling mechanical test process of the satellite and the solar wing can be referred to GB / T 34516-2017 for implementation.

[0054] The test process for installing the solar wing on the satellite body can be referred to GJB 2204A-2022 for implementation.

[0055] The specific test method can be referred to GB / T 34516-2017 for implementation.

[0056] It should be understood that although this specification is described according to each embodiment, not every embodiment only contains an independent technical solution. This narrative way of the specification is only for clarity. Those skilled in the art should regard the specification as a whole, and the technical solutions in each embodiment can also be appropriately combined to form other embodiments that can be understood by those skilled in the art.

[0057] The above are only the schematic specific embodiments of the present application, and are not used to limit the scope of the present application. Any equivalent changes, modifications and combinations made by those skilled in the art without departing from the concept and principles of the present application shall fall within the scope of protection of the present application.

Claims

1. A mass production satellite solar wing decoupling mechanical test method, comprising: S1: Establish a first satellite finite element model and a second satellite finite element model, wherein the first satellite finite element model is a whole satellite finite element model with solar panels, and the second satellite finite element model is a satellite body finite element model; S2: Modal analysis and sinusoidal response analysis are performed on the first satellite finite element model and the second satellite finite element model respectively to compare the differences in the stiffness simulation values ​​of the satellite body and the differences in the acceleration response simulation values ​​of the focus position between the two states of with and without solar wings; S3: Carry out three-directional sinusoidal characteristic level frequency sweep test on the satellite body to obtain the acceleration response of the satellite body's focus position; S4: Carry out three-directional sinusoidal characteristic level frequency sweep test on the solar wing unit to obtain the acceleration response of the focus position on the solar wing unit; S5: Satellite solar wing assembly; S6: Carry out a three-directional sinusoidal characteristic level frequency sweep test on the satellite body equipped with the solar wing, and obtain the acceleration response of the satellite body focus position that is the same as step S3, and the acceleration response of the solar wing stand-alone focus position that is the same as step S4; S7: Compare the acceleration response of the satellite body's focus position obtained in step S6 with the acceleration response of the satellite body's focus position obtained in step S3, so as to obtain the difference in acceleration response of the satellite body's focus position in the two states of having a solar wing and having no solar wing; S8: Compare the acceleration response of the solar wing stand-alone focus position obtained in step S6 with the acceleration response of the solar wing stand-alone focus position obtained in step S4, so as to obtain the difference in acceleration response of the solar wing stand-alone focus position under the two states of having a solar wing and having no solar wing; S9: Evaluate the decoupling consistency.

2. The method according to claim 1, wherein: In step S1, the first satellite finite element model is a finite element model established for a satellite body equipped with solar wings, and the second satellite finite element model is a finite element model established for a satellite body without solar wings.

3. The method according to claim 1, wherein: Step S2 includes, in finite element software, calculating the acceleration response simulation value of the focus position of the satellite body under acceptance-level sinusoidal vibration excitation in two states, with solar wings and without solar wings, by performing sinusoidal response analysis, and obtaining the difference in the acceleration response simulation values ​​of the satellite body in the two states, with solar wings and without solar wings.

4. The method according to claim 1, wherein: Step S2 includes, in finite element software, performing modal analysis to obtain the stiffness simulation value of the satellite body in two states, with solar wings and without solar wings, and obtaining the difference in stiffness simulation values ​​of the satellite body in two states, with solar wings and without solar wings.

5. The method according to claim 1, wherein: In step S4, the measuring points of the solar wing stand-alone sensor are not less than the measuring points of the solar wing in the whole-satellite mechanical test, and cover the measuring points of the solar wing in the whole-satellite mechanical test.

6. The method according to claim 1 or 3, wherein: The acceleration response includes the peak value and the frequency of occurrence of the peak value.

7. The method according to claim 1, wherein: In step S9, if the decoupling consistency evaluation result is poor, it is determined that the subsequent mass production satellite and solar wing decoupling mechanical test cannot be carried out.

8. The method according to claim 1, wherein: In step S9, if the decoupling consistency evaluation result is relatively poor, it is determined that the subsequent mass production satellite and solar wing decoupling mechanical test cannot be carried out; if the decoupling consistency evaluation result is relatively good, it is determined that the subsequent mass production satellite and solar wing decoupling mechanical test can be carried out.

9. The method according to claim 1, wherein: The evaluation indicators for evaluating the decoupling consistency in step S9 include: Step S2 obtains the difference in stiffness simulation value of the satellite body and the difference in acceleration response simulation value of the focus position under the two states of with solar wings and without solar wings; The difference in acceleration response of the satellite body's focus position between the two states of having solar panels and not having solar panels obtained in step S7; Step S8 obtains the difference in acceleration response of the solar wing unit's focus position in the two states of having a solar wing and not having a solar wing.

10. A computer-readable storage medium having stored thereon software instructions which, when executed, implement the method of any one of claims 1-9.

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