Clamp for random vibration test of spacecraft bolt assembly part and test method
By designing a random vibration test fixture for spacecraft bolt assemblies and conducting finite element analysis, the problem of inaccurate simulation of existing devices was solved, and accurate simulation of the real and complex vibration environment of spacecraft bolt assemblies was achieved, thereby improving the accuracy and reliability of the test.
Patent Information
- Application Number
- CN202510810835.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-06-17
- Publication Date
- 2025-09-16
AI Technical Summary
Existing random vibration test equipment cannot accurately simulate the performance of spacecraft bolt assemblies in real complex vibration environments, and the test methods fail to fully reflect the actual conditions of the spacecraft during launch and flight, resulting in insufficient accuracy in design applications.
A fixture for random vibration testing of spacecraft bolt assemblies was designed to simulate the common bolt connection method in spacecraft. Random vibration loads in different directions were applied by changing the clamping direction. Combined with the finite element analysis optimization model, the real complex vibration environment of spacecraft bolt assemblies was simulated.
It achieves accurate simulation of spacecraft bolt assemblies in real complex vibration environments, improves the accuracy and reliability of the test, and provides support for scientific research and engineering design.
Smart Images
Figure CN120645147A_ABST
Abstract
Description
Technical Field
[0001] The present invention relates to the technical field of spacecraft bolt assemblies, and in particular to a fixture and a test method for random vibration testing of spacecraft bolt assemblies. Background Art
[0002] Spacecraft bolt assemblies play a vital role in modern aerospace engineering. They withstand tremendous impact forces during launch and endure complex and varied dynamic loads and vibration environments. With the continuous advancement of aerospace technology, spacecraft designs are becoming increasingly complex. Bolted assemblies, due to their high reliability and ease of assembly, are widely used to connect various spacecraft components, ensuring the stability and functionality of the overall structure. However, these joints often become weak links in the structure, especially in random vibration environments, prone to damage and failure, affecting the safety and service life of the spacecraft.
[0003] To more accurately simulate and evaluate the performance of spacecraft bolt assemblies in random vibration environments, random vibration testing technology has been gradually introduced into the aerospace field in recent years. Existing random vibration testing equipment is mostly designed for general mechanical structures and has limited applicability to spacecraft-specific bolt assemblies, especially those with unique geometric angles and assembly requirements. Furthermore, the vibration environment experienced by spacecraft during launch and flight is extremely complex, including random vibrations of varying frequencies, amplitudes, and directions.
[0004] Traditional bolt assembly testing methods primarily rely on mechanical property testing under static or simplified dynamic loading conditions, failing to fully account for the impact of complex random vibrations in real-world spaceflight environments. While these testing devices can provide a preliminary assessment of the strength of spacecraft bolt assemblies, they struggle to fully reflect the complex vibration environments experienced by spacecraft during actual launch flights. This inability to accurately predict safety issues and failure modes in these complex vibration environments limits their application in spacecraft design.
[0005] Furthermore, random vibration testing of spacecraft bolt assemblies presents numerous challenges. For one thing, spacecraft components are typically large and complex. Furthermore, due to the high cost of developing spacecraft, the number of test specimens is limited. Therefore, consideration was given to simplifying the model calculations by approximating its scale, and developing a fixture and test method for random vibration testing of spacecraft bolt assemblies.
[0006] An ideal fixture for random vibration testing of spacecraft bolt assemblies should possess the following characteristics: it considers common bolt connection methods used in spacecraft; it can simulate the complex random vibration environment experienced by spacecraft during launch; it considers the inherent dynamic characteristics of spacecraft bolt assemblies, such as modal frequency and damping ratio; and it allows for approximate proportional simplification of the model calculations. Such a test fixture not only comprehensively considers the accuracy and reliability of spacecraft design but also provides strong experimental support for academic research in the field of aerospace engineering, possessing significant scientific value and engineering practical significance. Summary of the Invention
[0007] In view of the deficiencies and defects involved in the above-mentioned background technology, the purpose of the present invention is to provide a fixture and test method for random vibration testing of spacecraft bolt assemblies, which can simulate the unique bolt assemblies in spacecraft and the real complex vibration environment they experience, and perform random vibration tests on several common joint types in spacecraft under vibration loads in different directions.
[0008] In order to solve the above technical problems, the present invention adopts the following technical solutions: a fixture for random vibration testing of spacecraft bolt assemblies, comprising a spacecraft bolt assembly test piece group, a support, a pressure block, and a bolt; the support is fixedly connected to the vibration table, the spacecraft bolt assembly test piece group comprises a test piece one and a test piece two, the two test pieces are fixedly connected by a first bolt, and the pressure block is arranged on the fixed end of the spacecraft bolt assembly test piece group and is fixed to the support by a second bolt.
[0009] Furthermore, specimen one is provided with a support connection hole and a structural member connection hole, and specimen two is provided with a structural member connection hole; the structural member connection hole is used to connect specimen one and specimen two through bolts, and the support connection hole is used to connect specimen one and the support through bolts.
[0010] Furthermore, the diameter of the structural member connection hole on specimens 1 and 2 is d, and the center distance of the bolts when arranged is 3d, the end distance is 2d, and the edge distance is 1.5d.
[0011] Furthermore, there are four types of joints in Specimen 1 and Specimen 2, which are specifically divided into: lap joint, butt joint, 60° angle joint and 120° angle joint.
[0012] Furthermore, the structural component connection holes on the test piece 1 and the test piece 2 in the spacecraft bolt assembly test piece group of each joint type are divided into: single-row hole connection and double-row hole connection, which are used to realize the bolt connection between the test piece 1 and the test piece 2.
[0013] Furthermore, two through holes are provided at the lower end of the support for fixing to the vibration table. The direction of the external random vibration load is the vertical vibration of the support, thereby realizing the loading of the random vibration load generated by the vibration table surface onto the spacecraft bolt assembly test piece group.
[0014] Furthermore, two holes are provided on the top surface of the support for assembling the test piece group with spacecraft bolts, thereby conducting random vibration tests with thickness direction loads;
[0015] The support has vertical holes on the side for assembling the spacecraft bolt test piece group, and then conducting random vibration tests with width direction loads;
[0016] There are holes opened horizontally on the side of the support for the spacecraft bolt assembly test piece group, and then the random vibration test of the longitudinal load is carried out.
[0017] A test method for a fixture used for random vibration testing of spacecraft bolt assemblies, comprising the following steps:
[0018] Step 1: Select the connection method between the shell and the load-bearing frame in the target spacecraft, including overlap, butt joint, and corner joint with an angle of 60° and a corner joint with an angle of 120°; after simplifying and reducing the actual spacecraft shell plate to an approximate proportion, determine the diameter of the support connection hole and the structural component connection hole d, and determine the center distance of the structural component connection holes on specimens 1 and 2 to be 3d, the end distance to be 2d, and the edge distance to be 1.5d;
[0019] After the test piece design is completed, the type of test piece to be studied is selected and the test piece is assembled with the corresponding bolts;
[0020] Step 2: Based on the actual environment of the spacecraft, three clamping schemes are designed considering the random vibration of the test piece in the thickness, length and width directions;
[0021] Determine the test plan and perform clamping according to the plan;
[0022] Step 3: Pre-test before the test to verify system stability and preliminarily evaluate the specimen response;
[0023] Step 4: Import the power spectrum density curve of the basic excitation load determined in step 3 into the test vibration table, build a random vibration response test device according to the current specimen type, and start the random vibration response test; at the same time, the sensor on the vibration table collects the power spectrum density of the basic excitation load actually output by the vibration table in real time;
[0024] Step 5: Repeat the above steps, change the joint mode and clamping direction of the test piece, and conduct multiple tests. The sensors on the vibration table collect multiple sets of test measured data in real time.
[0025] Step 6: Based on the power spectrum density of the basic excitation load actually output by the vibration table during the test, perform structural random vibration analysis on the model again in the software to obtain data such as the response power spectrum density curve. Compare and analyze the data obtained from the software simulation with the actual test data to determine the accuracy of the finite element model, optimize the model, and improve the accuracy of the dynamic response.
[0026] Furthermore, the specific steps of step 3 are:
[0027] Step 3.1: Construct a finite element model of the test object and conduct structural modal analysis. Based on the first-order natural frequency of the structure, reasonably set the basic excitation load frequency range of random vibration;
[0028] Step 3.2: Determine the power spectrum density curve of the basic excitation load according to the frequency range of the random vibration load and conduct a random vibration simulation analysis of the structure.
[0029] In step 3.3, analyze the response power spectrum density curve obtained from the simulation. If the peak frequency of the curve is within the pre-set basic excitation frequency range, it indicates that the vibration characteristics of the test system meet the standards, and proceed to step 4. Otherwise, return to step 3.2 and readjust.
[0030] Compared with the prior art, the present invention adopts the above technical solution and has the following technical effects:
[0031] 1. Select bolt connection structure types commonly found in spacecraft components and simplify and reduce them to approximate proportions based on actual spacecraft shell panels. This avoids the limitations of existing bolt test pieces that only target common mechanical structures and restores the common bolt assemblies found in real spacecraft.
[0032] 2. By changing the clamping direction, random vibration of the specimen in three directions (thickness, length, and width) is achieved, achieving the effect of applying random vibration loads in different directions. This avoids the limitation of existing fixtures that can only test in a single direction, and better simulates the real complex vibration environment of spacecraft bolt assemblies.
[0033] 3. It is easy to understand, easy to operate, saves test time, and is universal for random vibration response test research of spacecraft bolt assemblies with different joint modes, materials and vibration environments. BRIEF DESCRIPTION OF THE DRAWINGS
[0034] Figure 1 Schematic diagram of the random vibration test clamping scheme of the spacecraft bolt assembly of the present invention;
[0035] Figure 2 Three random vibration test clamping schemes for single-row hole overlap: (a) thickness direction load vibration scheme, (b) length direction load vibration scheme, and (c) width direction load vibration scheme.
[0036] Figure 3 Schematic diagram of a spacecraft bolt structure; wherein, (a) is a schematic diagram of a single-row hole overlap structure, (b) is a schematic diagram of a double-row hole overlap structure, (c) is a schematic diagram of a single-row hole butt joint structure, (d) is a schematic diagram of a double-row hole butt joint structure, (e) is a schematic diagram of a single-row hole 60° angle joint structure, (f) is a schematic diagram of a double-row hole 60° angle joint structure, (g) is a schematic diagram of a single-row hole 120° angle joint structure, and (h) is a schematic diagram of a double-row hole 120° angle joint structure;
[0037] Figure 4 is a schematic diagram of the three-dimensional structure of the support 2;
[0038] Figure 5 3D structural diagram of the compact 3;
[0039] Figure 6 3D structural diagram of bolt 4, wherein (a) is a bolt 4-1 with a specification of M3, and (b) is a bolt 4-2 with a specification of M4;
[0040] Figure 7 This is a schematic diagram of the locations of the structural connection holes on specimens one and two.
[0041] Reference numerals: 1-spacecraft bolt assembly test piece, 2-support, 3-pressing block, 4-bolt; other test pieces connected with Figure 2 similar. DETAILED DESCRIPTION
[0042] The terms used in the present invention are only for the purpose of illustrating the embodiments of the present invention and are not intended to limit the present invention. Figure 1-7 , some embodiments of the present invention are described in detail. The present invention can be implemented in many different forms and should not be considered limited to the examples described herein. Rather, these examples are provided so that the present invention will be thorough and complete and will fully convey the scope of the invention to those skilled in the art. In the drawings, components are exaggerated for clarity.
[0043] like Figure 1 、 Figure 2 The figure shows a fixture for random vibration testing of spacecraft bolt assemblies, comprising a spacecraft bolt assembly test piece set 1, a support 2, a pressure block 3, and bolts 4. The joint method and clamping direction of the spacecraft bolt assembly test piece set 1 are adjustable design variables. During the test, the support 2 is fixedly connected to the vibration table. The spacecraft bolt assembly test piece set 1 comprises two structural components, Specimen 1 and Specimen 2, which are fixedly connected by a first bolt 4-1. The pressure block 3 is positioned on the fixed end of the spacecraft bolt assembly test piece set 1 and secured to the support 2 by a second bolt 4-2.
[0044] Specimen 1 is provided with support connection holes and structural member connection holes, and specimen 2 is provided with structural member connection holes; both are connected by bolts and are the main part of the test piece of the random vibration test fixture for spacecraft bolt assemblies;
[0045] Among them, the structural component connection holes on Specimens 1 and 2 in Spacecraft Bolt Assembly Test Specimen Group 1 are determined to have a commonly used diameter d of 3mm for side panel connection according to the spacecraft bolt connection hole diameter selection standard, and the center distance of the bolts is 3d, the end distance is 2d, and the edge distance is 1.5d;
[0046] The support connection hole on the test piece 1 in the spacecraft bolt assembly test piece group 1 is selected with a common diameter of 4mm;
[0047] like Figure 3 As shown in the figure, the specimens of the spacecraft bolt assembly test piece group 1 have four types of joints, namely lap joint, butt joint, 60° angle joint and 120° angle joint. These four types of joints are all taken from the common connection methods between the spacecraft shell and the load-bearing frame, and are simplified and reduced in approximate proportions according to the actual spacecraft shell plate.
[0048] The structural connection holes on Specimens 1 and 2 in Spacecraft Bolt Assembly Test Specimens 1, each with a joint type, are divided into single-row and double-row hole connections, used to bolt Specimens 1 and 2 together. Bolt connection structures commonly found in spacecraft components were selected to overcome the limitations of existing bolt test specimens, which only target common mechanical structures, and to replicate the bolt assembly components commonly found in real spacecraft.
[0049] like Figure 4 As shown, the support 2 is the fixed part of the random vibration test fixture of the spacecraft bolt assembly test piece, and its upper end is provided with a through hole for fixing the spacecraft bolt assembly test piece group 1 and the pressure block 3; the lower end of the support 2 is provided with two through holes for fixing to the vibration table, so as to realize the loading of the random vibration load generated by the vibration table surface onto the spacecraft bolt assembly test piece group 1.
[0050] Among them, the top surface of the support 2 is provided with two holes for assembling the test piece group 1 with spacecraft bolts, and then performing random vibration tests with thickness direction loads;
[0051] The side of the support 2 is provided with vertical holes for assembling the test piece group 1 with spacecraft bolts, thereby conducting random vibration tests with width-direction loads;
[0052] The side of the support 2 is provided with a hole for assembling the test piece group 1 with spacecraft bolts, thereby conducting a random vibration test of the longitudinal load;
[0053] Compared with traditional common random vibration test devices, the present invention achieves the effect of applying random vibration loads in different directions by changing the clamping direction, avoiding the limitation of existing fixtures that can only test in a single direction, and better simulating the real complex vibration environment of spacecraft bolt assemblies.
[0054] like Figure 5 As shown, the pressure block 3 is a rectangular parallelepiped, which is the fixed part of the random vibration test fixture of the spacecraft bolt assembly test piece. It is provided with a through hole for fixed connection, which is also 4mm. It is fixed on the support 2 together with the spacecraft bolt assembly test piece group 1 to prevent errors caused by different vibration starting positions in the random vibration test.
[0055] like Figure 1 、 Figure 2 As shown, after the spacecraft bolt assembly test piece and the pressure block 3 are fixed on the support 2 by the bolts 4-2, the random vibration response test of the spacecraft bolt assembly test piece is carried out. The direction of the external random vibration load is the vertical vibration of the support 2. As the clamping direction of the test piece changes, the vibration of the load occurs in three different directions: the length direction, the width direction, and the thickness direction of the test piece. By changing the joint method and the clamping direction of the test piece, the real complex vibration environment of the spacecraft bolt assembly is simulated, thereby making the vibration response of the test piece under the action of the random vibration load.
[0056] The present invention also discloses a fixture and a test method for random vibration testing of a spacecraft bolt assembly, comprising the following steps:
[0057] Step 1: Select common connection methods between the target spacecraft's outer shell and the load-bearing frame, including lap joints, butt joints, and corner joints with angles of 60° and 120°. After simplifying and reducing the actual spacecraft's outer shell panels to an approximate scale, and based on the spacecraft bolt connection hole diameter selection criteria, determine the commonly used holes with a diameter of 3mm for side panel connections. The center distance between the structural component connection holes on Specimens 1 and 2 is determined to be 3d, the end distance is 2d, and the edge distance is 1.5d. Once the test piece design is completed, determine the type of test piece to be studied, and assemble the test pieces using the appropriate bolts.
[0058] Step 2: Select different clamping methods based on the spacecraft's actual environment. Consider random vibrations in the thickness, length, and width of the test piece and design three clamping schemes. Determine the scheme to be tested and perform the clamping according to it.
[0059] Step 3: Pre-test before the test to verify system stability and preliminarily evaluate the specimen response;
[0060] Step 3.1: Construct a finite element model of the test object, conduct structural modal analysis, and reasonably set the basic excitation load frequency range of random vibration based on the first-order natural frequency of the structure.
[0061] Step 3.2: Determine the power spectrum density curve of the basic excitation load according to the frequency range of the random vibration load and carry out the random vibration simulation analysis of the structure.
[0062] In step 3.3, analyze the response power spectrum density curve obtained from the simulation. If the peak frequency of the curve is within the pre-set basic excitation frequency range, it indicates that the vibration characteristics of the test system meet the standards, and proceed to step 4. Otherwise, return to step 3.2 and readjust.
[0063] In step 4, the power spectrum density curve of the base excitation load determined in step 3 is imported into the test vibration table. The random vibration response test apparatus is set up according to the current specimen type, and the random vibration response test is started. Simultaneously, the sensors on the vibration table collect the power spectrum density of the base excitation load actually output by the vibration table in real time.
[0064] Step 5: Repeat the above steps, change the joint mode and clamping direction of the test piece to carry out multiple tests, and obtain multiple sets of test measured data in real time through the sensors on the vibration table.
[0065] Step 6: Based on the power spectrum density of the basic excitation load actually output by the vibration table during the test, perform structural random vibration analysis on the model again in the software to obtain data such as the response power spectrum density curve. Compare and analyze the data obtained from the software simulation with the actual test data to determine the accuracy of the finite element model, optimize the model, and improve the accuracy of the dynamic response.
[0066] This test device invention can provide benchmark data for calibrating the finite element dynamic analysis model of spacecraft bolt assemblies and improving the accuracy of dynamic response analysis by comparing the experimental measured data of random vibration response and the finite element analysis results.
[0067] This experimental invention can conduct experimental research on spacecraft bolt assemblies by changing the joint method and clamping direction, which is beneficial to avoid the limitations of existing random vibration fixtures and test plans, more accurately restore common bolt assemblies in spacecraft, simulate the real and complex vibration environment of spacecraft bolt assemblies, and provide benchmark data for verifying the accuracy of finite element dynamic analysis models and dynamic response analysis of spacecraft bolt assemblies.
[0068] The experimental device invention is universal for random vibration test research of spacecraft bolt assemblies with different joint modes, different materials and different environments, is easy to understand, easy to operate, and saves test time.
[0069] It will be understood by those skilled in the art that, unless otherwise defined, all terms (including technical and scientific terms) used herein have the same meaning as commonly understood by those skilled in the art in the art to which the present invention belongs. It should also be understood that terms such as those defined in common dictionaries should be understood to have meanings consistent with their meanings in the context of the prior art and, unless defined as such, will not be interpreted in an idealized or overly formal sense.
[0070] The specific implementation methods described above further illustrate the objectives, technical solutions and beneficial effects of the present invention in detail. It should be understood that the above description is only a specific implementation method of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.
Claims
1. A fixture for random vibration testing of spacecraft bolt assemblies, characterized in that: It includes a spacecraft bolt assembly test piece group, a support, a pressure block, and a bolt; the support is fixedly connected to the vibration table, the spacecraft bolt assembly test piece group includes test piece one and test piece two, the two test pieces are fixedly connected by a first bolt, and the pressure block is arranged on the fixed end of the spacecraft bolt assembly test piece group and is fixed to the support by a second bolt.
2. The fixture for random vibration testing of spacecraft bolt assemblies according to claim 1, characterized in that: Specimen one is provided with support connection holes and structural member connection holes, and specimen two is provided with structural member connection holes; the structural member connection holes are used to connect specimen one and specimen two through bolts, and the support connection holes are used to connect specimen one and the support through bolts.
3. The fixture for random vibration testing of spacecraft bolt assemblies according to claim 2, characterized in that: The diameter of the structural member connection hole on specimens 1 and 2 is d, and the center distance of the bolts when arranged is 3d, the end distance is 2d, and the edge distance is 1.5d.
4. The fixture for random vibration testing of spacecraft bolt assemblies according to claim 1, characterized in that: There are four types of joints in specimens 1 and 2, which are specifically divided into: lap joint, butt joint, 60° angle joint and 120° angle joint.
5. The fixture for random vibration testing of spacecraft bolt assemblies according to claim 4, characterized in that: The structural component connection holes on the test piece one and the test piece two in the spacecraft bolt assembly test piece group of each joint form are further divided into: single-row hole connection and double-row hole connection, which are used to realize the bolt connection between the test piece one and the test piece two.
6. The fixture for random vibration testing of spacecraft bolt assemblies according to claim 1, characterized in that: Two through holes are provided at the lower end of the support for fixing to the vibration table. The direction of the external random vibration load is the vertical vibration of the support, so that the random vibration load generated by the vibration table surface is loaded onto the spacecraft bolt assembly test piece group.
7. The fixture for random vibration testing of spacecraft bolt assemblies according to claim 1, characterized in that: There are two holes on the top surface of the support for the spacecraft bolt assembly test piece group, and then the random vibration test of the thickness direction load is carried out; The support has vertical holes on the side for assembling the spacecraft bolt test piece group, and then conducting random vibration tests with width direction loads; There are holes opened horizontally on the side of the support for the spacecraft bolt assembly test piece group, and then the random vibration test of the longitudinal load is carried out.
8. A test method for a fixture used for random vibration testing of spacecraft bolt assemblies, characterized in that: The specific steps are: Step 1: Select the connection method between the shell and the load-bearing frame in the target spacecraft, including overlap, butt joint, and corner joint with an angle of 60° and a corner joint with an angle of 120°; after simplifying and reducing the actual spacecraft shell plate to an approximate proportion, determine the diameter of the support connection hole and the structural component connection hole d, and determine the center distance of the structural component connection holes on specimens 1 and 2 to be 3d, the end distance to be 2d, and the edge distance to be 1.5d; After the test piece design is completed, the type of test piece to be studied is selected and the test piece is assembled with the corresponding bolts; Step 2: Based on the actual environment of the spacecraft, three clamping schemes are designed considering the random vibration of the test piece in the thickness, length and width directions; Determine the test plan and perform clamping according to the plan; Step 3: Pre-test before the test to verify system stability and preliminarily evaluate the specimen response; Step 4: Import the power spectrum density curve of the basic excitation load determined in step 3 into the test vibration table, build a random vibration response test device according to the current specimen type, and start the random vibration response test; at the same time, the sensor on the vibration table collects the power spectrum density of the basic excitation load actually output by the vibration table in real time; Step 5: Repeat the above steps, change the joint mode and clamping direction of the test piece, and conduct multiple tests. The sensors on the vibration table collect multiple sets of test measured data in real time. Step 6: Based on the power spectrum density of the basic excitation load actually output by the vibration table during the test, perform structural random vibration analysis on the model again in the software to obtain data such as the response power spectrum density curve. Compare and analyze the data obtained from the software simulation with the actual test data to determine the accuracy of the finite element model, optimize the model, and improve the accuracy of the dynamic response.
9. The test method of the fixture for random vibration testing of spacecraft bolt assemblies according to claim 8, characterized in that: The specific steps of step 3 are: Step 3.1: Construct a finite element model of the test object and conduct structural modal analysis. Based on the first-order natural frequency of the structure, reasonably set the basic excitation load frequency range of random vibration; Step 3.2: Determine the power spectrum density curve of the basic excitation load according to the frequency range of the random vibration load and conduct a random vibration simulation analysis of the structure. In step 3.3, analyze the response power spectrum density curve obtained from the simulation. If the peak frequency of the curve is within the pre-set basic excitation frequency range, it indicates that the vibration characteristics of the test system meet the standards, and proceed to step 4. Otherwise, return to step 3.2 and readjust.