Batch production satellite solar wing decoupling mechanical test method

Abstract

The application provides a batch satellite solar wing decoupling mechanical test method, comprising: establishing a satellite finite element model and comparing the responses of two states of with and without solar wings, then performing a mechanical test on the satellite alone; performing a mechanical test on the solar wing alone; performing a mechanical test on the solar wing after being installed on the satellite; evaluating the responses of the solar wing in the two states of with and without solar wings and the response of the satellite; evaluating the decoupling results; and performing or terminating the batch satellite solar wing decoupling mechanical test implementation according to the evaluation results. The method can not only meet the vibration test examination of the satellite body and the solar wing, but also reduce the number of times of installing the solar wing on the satellite, and improve the development efficiency of the batch satellite.

Description

TECHNICAL FIELD

[0001] The present application relates to a satellite dynamics test method, in particular to a batch satellite solar wing decoupling dynamics test method. BACKGROUND

[0002] The state of satellite test technology is difficult to obtain, but from the satellite dynamics test diagram published to the outside world, it can be seen that the satellite is a solar wing dynamics test. At present, the satellite development is all carried out with solar wing dynamics test. For example, the foreign standard "Test requirements for launch, upper-stage and space vehicles" (SMC-S-016) stipulates that all flight products and software should be completed before the system level test. The system level test project, purpose and related test requirements are stipulated in the national military standard "Test requirements for launch, upper-stage and space vehicles" (GJB 1027A-2020), and the state of the spacecraft described in the vibration test description is the launch state.

[0003] Batch satellites have the characteristics of high overall state similarity, short development cycle, fast launch requirement, etc., which puts forward high requirements on satellite development process and development efficiency. However, in the traditional satellite development process, the solar wing installation and electric explosion process before and after the single satellite dynamics test is long, involves many posts, and consumes a lot of manpower and material resources. SUMMARY

[0004] In view of the deficiencies in the prior art, for batch satellite launch tasks, the present application provides a batch satellite solar wing decoupling dynamics test method, which comprehensively designs the comprehensiveness of the process and the coverage and reliability of the test process, can effectively meet the satellite dynamics characteristics examination and screening, optimize the number of solar wing installation and disassembly, improve the utilization efficiency of personnel and solar wing deployment equipment resources, and meet the needs of batch satellite.

[0005] The present application provides a batch satellite solar wing decoupling dynamics test method, comprising:

[0006] S1: establishing 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 wing, and the second satellite finite element model is a satellite body finite element model;

[0007] S2: modal analysis and sinusoidal response analysis are carried out on the first satellite finite element model and the second satellite finite element model respectively, so as to compare the stiffness simulation value difference and the acceleration response simulation value difference of the satellite body in the two states of with solar wing and without solar wing;

[0008] S3: Conduct a three-directional sinusoidal characteristic-level frequency sweep test on the satellite body to obtain the acceleration response of the satellite body at the location of interest;

[0009] S4: Conduct a three-directional sinusoidal characteristic level frequency sweep test on a single solar array to obtain the acceleration response at the location of interest on the single solar array;

[0010] S5: Satellite Solar Winged Star;

[0011] S6: Conduct a three-directional sinusoidal characteristic level frequency sweep test on the satellite body equipped with solar arrays to obtain the acceleration response of the satellite body at the location of interest, which is the same as in step S3, and the acceleration response of the solar array at the location of interest, which is the same as in step S4.

[0012] S7: Compare the acceleration response of the satellite body at the point of interest obtained in step S6 with the acceleration response of the satellite body at the point of interest obtained in step S3, so as to obtain the difference in acceleration response of the satellite body at the point of interest under the two states of having solar panels and not having solar panels.

[0013] S8: Compare the acceleration response of the solar array single-unit focus position obtained in step S6 with the acceleration response of the solar array single-unit focus position obtained in step S4, so as to obtain the difference in acceleration response of the solar array single-unit focus position in the two states of having a solar array and not having a solar array.

[0014] S9: Evaluate the consistency of decoupling.

[0015] According to an embodiment of the method provided in this application, in step S1, the first satellite finite element model is a finite element model established for a satellite body equipped with solar panels, and the second satellite finite element model is a finite element model established for a satellite body that does not contain solar panels.

[0016] According to an embodiment of the method provided in this application, step S2 includes, in finite element software, performing sinusoidal response analysis to calculate the simulated acceleration response values ​​of the satellite body's focus position under acceptance-level sinusoidal vibration excitation in two states: with and without solar panels, and obtaining the difference in the simulated acceleration response values ​​of the satellite body in the two states: with and without solar panels.

[0017] According to an embodiment of the method provided in this application, step S2 includes, in finite element software, performing modal analysis to obtain the stiffness simulation values ​​of the satellite body in two states: with and without solar panels, and obtaining the difference between the stiffness simulation values ​​of the satellite body in the two states: with and without solar panels.

[0018] According to an embodiment of the method provided in this application, in step S4, the number of sensor measurement points on the solar array is not less than the number of measurement points on the solar array in the whole-satellite mechanics test, and covers the number of measurement points on the solar array in the whole-satellite mechanics test.

[0019] According to an embodiment of this application, the acceleration response includes a peak value and the frequency at which the peak value occurs.

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

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

[0022] According to an embodiment of the method provided in this application, the evaluation metrics for assessing decoupling consistency in step S9 include:

[0023] Step S2 yields the differences in simulated stiffness values ​​of the satellite body with and without solar panels, as well as the differences in simulated acceleration response values ​​at the points of interest.

[0024] Step S7 yields the difference in acceleration response at the target location of the satellite body under two states: with and without solar panels.

[0025] Step S8 yields the difference in acceleration response at the point of interest for a single solar array unit under two states: with and without solar arrays.

[0026] This application also provides a computer-readable storage medium having software instructions stored thereon, which, when executed, implement the method described above.

[0027] This method can not only meet the vibration test requirements of the satellite body and solar array, but also reduce the number of times the solar array is installed on mass-produced satellites, thereby improving the development efficiency of mass-produced satellites.

[0028] The advantages of this invention compared to the prior art are as follows:

[0029] ①This invention provides a mechanical testing method for mass-produced satellites without solar panels.

[0030] ②The test method provided by this invention optimizes the number of times the solar array is installed and disassembled, improves the efficiency of personnel and solar array deployment equipment resource utilization, and reduces human and material costs.

[0031] ③The testing method of this invention shortens the R&D cycle of mass-produced satellites and increases the launch speed of mass-produced satellites, and has a wide range of application and promotion value. Attached Figure Description

[0032] The following description, in conjunction with the accompanying drawings, will further illustrate the above-mentioned features, technical characteristics, advantages, and implementation methods of this application in a clear and understandable manner. The accompanying drawings are for illustrative and explanatory purposes only and do not limit the scope of this application. Wherein:

[0033] Figure 1 This is a flowchart of the decoupling mechanics test method for mass-produced satellite solar panels according to the present invention.

[0034] Figure 2 The process of developing a mass-produced satellite and solar array decoupling mechanics test is shown in the prior art. Detailed Implementation

[0035] To provide a clearer understanding of the technical features, objectives, and effects of this application, specific embodiments of this application will now be described with reference to the accompanying drawings.

[0036] This application provides a method for mechanical testing of mass-produced satellites without solar panels, optimizing the testing process before and after the experiment. Figure 2 As shown, compared to the traditional satellite development process, the optimized process involves the installation, manual deployment, and illumination testing of the satellite and solar array before the mechanical test; the installation, testing, and clamping of the pyrotechnic device; and the electric explosion deployment, illumination testing, and removal of the solar array after the mechanical test. During the launch-ready modification, all of the above-mentioned work items for the solar array are tested, and the actual power supply capability of the pyrotechnic device is tested after the mechanical test and before this stage.

[0037] One embodiment of this application provides a decoupling mechanics test method for mass-produced satellite solar panels, including the following steps:

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

[0039] S2: Modal analysis and sinusoidal response analysis were performed on the finite element model of the first satellite and the finite element model of the second satellite respectively, in order to compare the differences in the simulated stiffness values ​​of the satellite body and the differences in the simulated acceleration response values ​​at the points of interest under the two states of having and not having solar panels.

[0040] Specifically, in finite element software, modal analysis is performed to obtain the simulated stiffness values ​​of the satellite body with and without solar panels, and the difference between the simulated stiffness values ​​of the satellite body with and without solar panels is obtained.

[0041] Specifically, in finite element analysis software, sinusoidal response analysis was conducted to calculate the simulated acceleration response (including peak value and frequency) of the satellite's target location under acceptance-level sinusoidal vibration excitation in both states with and without solar panels. The differences in the simulated acceleration response values ​​between the two states were then obtained. If the peak value and frequency of the satellite's response are approximately the same, it indicates that the presence or absence of solar panels has minimal impact on the satellite's mechanical transmission characteristics, and the risk is controllable.

[0042] S3: Conduct a three-directional sinusoidal characteristic level frequency sweep test on the satellite body (excluding the solar panels) to obtain the acceleration response of the satellite body at the location of interest. The specific test method can be referred to QJ 1579A-2005.

[0043] S4: Conduct a three-directional sinusoidal characteristic level frequency sweep test on a single solar array unit. The number of sensor measurement points on the single solar array unit shall be no less than the number of measurement points on the solar array in the whole satellite mechanics test, and shall cover the number of measurement points on the solar array in the whole satellite mechanics test, to obtain the acceleration response of the location of interest on the single solar array unit.

[0044] S5: Install satellite solar panels and confirm that the sensors on the satellite and solar panels are installed correctly;

[0045] S6: After the satellite is equipped with solar panels, a three-directional sinusoidal characteristic level frequency sweep test is carried out on the entire satellite (i.e., the satellite body equipped with solar panels) to obtain the acceleration response of the satellite body at the point of interest, which is the same as in step S3, and the acceleration response of the solar panel at the point of interest, which is the same as in step S4.

[0046] S7: Compare the acceleration response of the satellite body at the point of interest obtained in step S6 with the acceleration response of the satellite body at the point of interest obtained in step S3, so as to obtain the difference in acceleration response of the satellite body at the point of interest with and without solar panels.

[0047] S8: Compare the acceleration response of the solar array single-unit focus position obtained in step S6 with the acceleration response of the solar array single-unit focus position obtained in step S4 to obtain the difference in acceleration response of the solar array single-unit focus position with and without the solar array.

[0048] S9: Evaluate the decoupling consistency. If the evaluation result is good, it is determined that subsequent decoupling mechanical tests of mass-produced satellites and solar panels can be carried out. The specific test method can refer to QJ 1579A-2005. If the evaluation result is poor, it is determined that subsequent decoupling mechanical tests of mass-produced satellites and solar panels cannot be carried out. The evaluation indicators include:

[0049] The differences in the simulated stiffness of the satellite body and the differences in the simulated acceleration response at the point of interest are obtained in step S2 under the two states with and without solar panels.

[0050] The difference in acceleration response of the satellite body at the point of interest is obtained in step S7 under the two states of having and not having solar panels;

[0051] Step S8 yields the difference in acceleration response at the point of interest for a single solar array unit under the two states of having and not having a solar array.

[0052] The above evaluation indicators can be used to determine whether the mechanical test response of a single solar array unit encompasses the magnitude of the whole satellite test, and whether the acceleration response of the un-solar array unit at the location of interest of the whole satellite can cover the whole satellite test state and meet the launch and single-unit test conditions.

[0053] The decoupling mechanical test procedure for satellite and solar array can be performed in accordance with GB / T 34516-2017.

[0054] The verification process for installing solar panels on the satellite body can be carried out in accordance with GJB 2204A-2022.

[0055] The specific test methods can be performed in accordance with GB / T 34516-2017.

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

[0057] The above description is merely an illustrative embodiment of this application and is not intended to limit the scope of this application. Any equivalent changes, modifications, and combinations made by those skilled in the art without departing from the concept and principles of this application shall fall within the scope of protection of this application.

Claims

1. A method for decoupling mechanical testing of solar panels in mass-produced satellites, comprising: S1: Establish the first satellite finite element model and the second satellite finite element model, where the first satellite finite element model is the whole satellite finite element model with solar panels, and the second satellite finite element model is the satellite body finite element model; S2: Modal analysis and sinusoidal response analysis were performed on the finite element model of the first satellite and the finite element model of the second satellite respectively, in order to compare the differences in the simulated stiffness values ​​of the satellite body and the differences in the simulated acceleration response values ​​at the points of interest under the two states of having solar panels and not having solar panels. S3: Conduct a three-directional sinusoidal characteristic-level frequency sweep test on the satellite body to obtain the acceleration response of the satellite body at the location of interest; S4: Conduct a three-directional sinusoidal characteristic level frequency sweep test on a single solar array to obtain the acceleration response at the location of interest on the single solar array; S5: Satellite Solar Winged Star; S6: Conduct a three-directional sinusoidal characteristic level frequency sweep test on the satellite body equipped with solar arrays to obtain the acceleration response of the satellite body at the location of interest, which is the same as in step S3, and the acceleration response of the solar array at the location of interest, which is the same as in step S4. S7: Compare the acceleration response of the satellite body at the point of interest obtained in step S6 with the acceleration response of the satellite body at the point of interest obtained in step S3, so as to obtain the difference in acceleration response of the satellite body at the point of interest under the two states of having solar panels and not having solar panels. S8: Compare the acceleration response of the solar array single-unit focus position obtained in step S6 with the acceleration response of the solar array single-unit focus position obtained in step S4, so as to obtain the difference in acceleration response of the solar array single-unit focus position in the two states of having a solar array and not having a solar array. S9: Evaluate the consistency of decoupling.

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

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

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

5. The method according to claim 1, wherein, In step S4, the number of sensor measurement points on the solar array unit is no less than the number of measurement points on the solar array in the whole-satellite mechanics test, and covers the number of measurement points on the solar array in the whole-satellite mechanics test.

6. The method according to claim 1 or 3, wherein, Acceleration response includes the peak value and the frequency 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 decoupling mechanical test between the mass-produced satellite and the solar array cannot be carried out subsequently.

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 decoupling mechanical test between the mass-produced satellite and the solar array cannot be carried out subsequently; if the decoupling consistency evaluation result is relatively good, it is determined that the decoupling mechanical test between the mass-produced satellite and the solar array can be carried out subsequently.

9. The method according to claim 1, wherein, The evaluation metrics for assessing decoupling consistency in step S9 include: Step S2 yields the differences in simulated stiffness values ​​of the satellite body with and without solar panels, as well as the differences in simulated acceleration response values ​​at the points of interest. Step S7 yields the difference in acceleration response at the target location of the satellite body under two states: with and without solar panels. Step S8 yields the difference in acceleration response at the point of interest for a single solar array unit under two states: with and without solar arrays.

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

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