Spacecraft triaxial sinusoidal vibration test system and test method

Abstract

The embodiment of the invention provides a spacecraft three-axis sinusoidal vibration test system and method, and the system is characterized in that the system comprises a main working platform, a connecting assembly, a plurality of vibration tables, a cooling sub-module, an oil pump sub-module, a power amplifier sub-module, and an MIMO vibration control sub-module; the spacecraft is connected with the main working platform through the connecting assembly, the spacecraft is located above the multiple three-way force measuring sensors, and the multiple vibration tables are controlled to conduct full-magnitude vibration tests based on preset requirements on the basis of preset three-axis sine test conditions so as to test relevant parameters of the spacecraft. The embodiment of the invention ensures that the product is fully and widely excited from all angles of the three-dimensional space, so that the space motion of the spacecraft fills the whole three-dimensional physical space, the track is richer, the assessment is more effective, and compared with the traditional uniaxial test method, the mechanical test efficiency and the batch production development efficiency of the spacecraft product are greatly improved.

Description

Technical Field

[0001] The embodiments of this specification relate to the technical field of spacecraft dynamics testing, and particularly to a spacecraft three-axis sine vibration test system and a test method. Background Art

[0002] The spacecraft three-axis vibration test technology is an advanced vibration test technology with simultaneous loading in three orthogonal directions. If the unidirectional test loads of the original single-axis vibration are simultaneously applied in each direction during multi-axis vibration, the loads will be superimposed on each other, resulting in excessive loads on the test product, causing serious over-testing and even product damage. Therefore, it is necessary to re-design the three-axis vibration test conditions. The design of three-axis vibration test conditions is one of the difficult problems in the development of three-axis test technology, mainly including the design of amplitude magnitude conditions for each axis and the design of phase relationship conditions between each axis.

[0003] The method for designing three-axis six-degree-of-freedom vibration test conditions based on interface force equivalence (CN202210258221.3) mainly starts from the resultant force and resultant moment received by the spacecraft-star rocket connection interface, and designs the amplitude magnitude conditions of the three-axis six-degree-of-freedom vibration test from the perspective of single / triaxial force equivalence. The method for trimming multi-axis vibration test conditions based on acceleration equivalence (CN201510797190.9) focuses on the acceleration response of key parts on the product. To ensure that the response in each direction during multi-axis vibration is the same as the response when the single axis is applied separately, a method for trimming the amplitude magnitude conditions of the multi-axis vibration test based on acceleration equivalence is designed with the single-axis vibration test conditions as the benchmark. The method for trimming multi-axis vibration test conditions based on fatigue damage equivalent analysis (CN201611047327.X) mainly starts from fatigue damage equivalence, which can ensure that the actual stress fatigue situation can be truly simulated during the multi-axis vibration test, so as to simulate the real test environment. At the same time, the positions on the product that are difficult to assess in the actual environment can be assessed through this method. Summary of the Invention

[0004] In view of this, the embodiments of this specification provide a spacecraft three-axis sine vibration test method. One or more embodiments of this specification simultaneously relate to a device for the spacecraft three-axis sine vibration test method, a computing device, a computer-readable storage medium, and a computer program to solve the technical defects existing in the prior art.

[0005] According to the first aspect of the embodiments of this specification, a spacecraft three-axis sine vibration test system is provided, including: a main working platform, a connection component, a plurality of vibration tables, a cooling sub-module, an oil pump sub-module, a power amplifier sub-module, and a MIMO vibration control sub-module, wherein,

[0006] The cooling sub-module is used to maintain the working temperatures of the main working platform and the power amplifier sub-module; the oil pump sub-module is used to provide a high-pressure oil source for the module and precisely control the vibration amplitude by adjusting the oil pressure; the power amplifier sub-module is used to amplify the low-power signal output by the MIMO control sub-module to the high-power level required by the main working platform; the MIMO vibration control sub-module is used to generate the excitation signals of each vibration table in real time based on the multi-input multi-output control algorithm;

[0007] A plurality of vibration tables are respectively arranged on one side or both sides of the main working platform in the first, second or third direction, and are used to provide translational and / or rotational motions in the first, second or third direction for the main working platform, wherein the first direction is a preset direction in the horizontal plane, the second direction is a direction perpendicular to the first direction in the horizontal plane, and the third direction is a direction perpendicular to the horizontal plane;

[0008] The connection assembly is arranged on the upper surface of the main working platform and is used to connect the main working platform and the spacecraft to be tested;

[0009] A plurality of three-axis force sensors are arranged on the connection assembly. Among them, the set X-axis, set Y-axis and set Z-axis of each set three-axis force sensor in the set three axes are respectively parallel to each other, and the set X-axis is parallel to the first direction, the set Y-axis is parallel to the second direction, and the set Z-axis is parallel to the third direction;

[0010] The spacecraft is connected to the main working platform through the connection assembly, the spacecraft is located above the plurality of three-axis force sensors, and a full-scale vibration test is controlled for the plurality of vibration tables based on the preset requirements according to the preset three-axis sine test conditions to test the relevant parameters of the spacecraft.

[0011] In some embodiments, the three-axis sine test conditions include:

[0012] The frequency value of each frequency point selected from M frequency points according to the set reference angle AN from the set frequency range [f s , f p is determined according to a preset first calculation strategy, wherein the first calculation strategy is related to the number of octaves, M = (360 / AN) 2 -1, and each frequency point is provided with a corresponding positive integer sequence number m, m ∈ [1, M];

[0013] The cyclic phase difference of the cyclic axis corresponding to each frequency point is calculated according to a preset second calculation strategy, and the distribution of the cyclic phase differences corresponding to all frequency points satisfies a sine distribution, wherein the cyclic axis is one of the set X-axis, set Y-axis and set Z-axis, and the cyclic phase difference is the phase difference of the cyclic axis corresponding to any frequency point with respect to the standard phase;

[0014] The growth phase difference of the growth axis corresponding to each frequency point is calculated according to a preset third calculation strategy, where the growth axis is one of the non-cyclic axes among the set X-axis, the set Y-axis, and the set Z-axis;

[0015] The sum of the phase differences between any two of the three axes corresponding to each frequency point is 0, where the phase difference between axes is the phase difference between any two of the set three axes.

[0016] In some embodiments, the first calculation strategy includes:

[0017] Combined with a preset first calculation formula, calculate the frequency value of each frequency point, where the first calculation formula includes:

[0018]

[0019] where N represents the number of octaves, f m represents the frequency value of the mth frequency point, represents the integer quotient of the dividend divided by the divisor, and "dividend" mod "divisor" represents the integer remainder obtained by dividing the dividend by the divisor.

[0020] In some embodiments, the second calculation strategy includes:

[0021] Combined with a preset second calculation formula, calculate the cyclic phase difference of each frequency point, where the second calculation formula includes:

[0022]

[0023] where, represents the cyclic phase difference corresponding to the mth frequency point.

[0024] In some embodiments, the third calculation strategy includes:

[0025] Combined with a preset third calculation formula, calculate the growth phase difference of each frequency point, where the third calculation formula includes:

[0026]

[0027] where, represents the growth phase difference corresponding to the mth frequency point.

[0028] In some embodiments, the three-axis sine test conditions further include:

[0029] The amplitude magnitude conditions of each axis in the set three axes are:

[0030] Based on interface force equivalence, and / or

[0031] Based on acceleration equivalence, and / or

[0032] Based on fatigue damage equivalence.

[0033] In some embodiments, a plurality of three - axis force sensors are connected to a preset signal conditioner. The signal conditioner is provided with a plurality of signal channels. Among them,

[0034] p signal channels of the signal conditioner are respectively connected to the set Z - axis of each three - axis force sensor, where the number of three - axis force sensors is p;

[0035] Another p signal channels of the signal conditioner are respectively connected to the combined signals of the set X - axis or Y - axis of any two three - axis force sensors.

[0036] In some embodiments, the number of vibration tables is 8, where,

[0037] Determine the first axis, the second axis, and the third axis that pass through the geometric centroid of the main working platform and are respectively parallel to the first direction, the second direction, and the third direction;

[0038] One vibration table with a vibration direction parallel to the first axis is arranged on each side of the first axis;

[0039] One vibration table with a vibration direction parallel to the second axis is arranged on each side of the second axis;

[0040] Four vibration tables that are arranged in a pairwise - symmetric rectangular pattern around the third axis and have a vibration direction parallel to the third axis are arranged.

[0041] In some embodiments, a plurality of measurement points are arranged on the surface and / or inside of the spacecraft. The measurement points are used to collect the vibration response data of each part of the spacecraft, evaluate the structural dynamic characteristics, and check for abnormal responses to assist in fault diagnosis.

[0042] According to the second aspect of the embodiments of this specification, a method for a spacecraft three - axis sine vibration test is provided. The method is applied to the aforementioned spacecraft three - axis sine vibration test system. The method includes:

[0043] Sequentially turn on the main vibration table, a plurality of vibration tables, the cooling sub - module, the oil pump sub - module, the power amplifier sub - module, and the MIMO vibration control sub - module of the system, check whether the operating states of all components are normal, and set the main vibration table to the coarse - middle position;

[0044] Use a sling to lift the spacecraft to be tested, position the spacecraft above the connection component, slowly lower it, and then fixedly connect the spacecraft to the connection component, and remove the sling. Among them, the spacecraft to be tested is initially set to the power - off state;

[0045] Carry out precise centering debugging on the main vibration table;

[0046] Turn on multiple three - axis force sensors, perform a pre - vibration - level conduction vibration test based on a preset first parameter, after self - inspection, process and record the data through a preset data - processing module, and detect the corresponding data of each measuring point and perform anomaly handling after the pre - vibration - level conduction vibration test ends;

[0047] Set the spacecraft under test to the power - on state, turn on multiple three - axis force sensors, perform the first characteristic - level vibration test based on a preset second parameter, after self - inspection, process and record the data through the data - processing module, monitor the working state of the spacecraft under test throughout the test process, and detect the corresponding data of each measuring point and perform anomaly handling after the pre - vibration - level conduction vibration test ends;

[0048] Turn on multiple three - axis force sensors, perform a full - scale vibration test based on preset three - axis sine test conditions, after self - inspection, process and record the data through the data - processing module, monitor the working state of the spacecraft under test throughout the test process, and evaluate the effectiveness of the three - axis sine test conditions after the full - scale vibration test ends;

[0049] Set the spacecraft under test to the power - on state, turn on multiple three - axis force sensors, perform the second characteristic - level vibration test based on a preset third parameter, after self - inspection, process and record the data through the data - processing module, monitor the working state of the spacecraft under test throughout the test process, compare the relevant data in the two characteristic - level vibration tests, judge whether the test results of the two characteristic - level vibration tests are consistent, and check whether there is any damage or other anomalies on the appearance of the spacecraft under test to determine whether the spacecraft under test can safely withstand the sine vibration test;

[0050] Dismantle the spacecraft under test and lift it off, turn off the main vibration table, multiple vibration tables, cooling sub - module, oil - pump sub - module, power - amplifier sub - module, and MIMO vibration control sub - module in sequence, and the test ends.

[0051] At least one embodiment in the embodiments of this specification conducts experiments on the above - mentioned spacecraft three - axis sine vibration test system according to preset three - axis sine test conditions, aiming to solve the problem that the spatial motion trajectory of the spacecraft is relatively single under the excitation of conventional three - axis sine test conditions, ensure that the product is fully and widely excited from all angles in three - dimensional space, make the spatial motion of the spacecraft fill the entire three - dimensional physical space, with a richer trajectory and more effective assessment, and compared with the traditional single - axis test method, it will greatly improve the mechanical test efficiency and mass - production research and development efficiency of spacecraft products. Brief Description of the Drawings

[0052] Figure 1 It is a simple schematic diagram of the setting position of three - axis force sensors in a spacecraft three - axis sine vibration test system provided by some embodiments of this specification;

[0053] Figure 2 It is a schematic diagram of the X-axis phase reference test conditions of a three-axis sine vibration test method for a spacecraft provided by some embodiments of this specification;

[0054] Figure 3 It is a schematic diagram of the Y-axis phase reference test conditions of a three-axis sine vibration test method for a spacecraft provided by some embodiments of this specification;

[0055] Figure 4 It is a schematic diagram of the Z-axis phase reference test conditions of a three-axis sine vibration test method for a spacecraft provided by some embodiments of this specification. Specific implementation manners

[0056] Many specific details are set forth in the following description in order to provide a thorough understanding of this specification. However, this specification can be implemented in many other ways different from those described herein, and those skilled in the art can make similar generalizations without departing from the connotation of this specification. Therefore, this specification is not limited by the specific implementations disclosed below.

[0057] The terms used in one or more embodiments of this specification are for the purpose of describing specific embodiments only and are not intended to limit one or more embodiments of this specification. The singular forms "a" and "the" used in one or more embodiments of this specification and the appended claims are also intended to include the plural forms unless the context clearly dictates otherwise. It should also be understood that the term "and / or" used in one or more embodiments of this specification refers to and encompasses any and all possible combinations of one or more of the associated listed items. The modifiers "a" and "multiple" mentioned in this disclosure are illustrative rather than restrictive, and those skilled in the art should understand that unless clearly stated otherwise in the context, it should be understood as "one or more".

[0058] It should be understood that although the terms first, second, etc. may be used in one or more embodiments of this specification to describe various information, such information should not be limited to these terms. These terms are only used to distinguish the same type of information from each other. For example, without departing from the scope of one or more embodiments of this specification, the first may also be referred to as the second, and similarly, the second may also be referred to as the first. Depending on the context, the word "if" as used herein can be interpreted as "when" or "while" or "in response to determining".

[0059] First, the noun terms related to one or more embodiments of this specification are explained.

[0060] MIMO: Multiple-Input Multiple-Output, which is multi-input multi-output in Chinese.

[0061] Some embodiments of this specification provide a three-axis sinusoidal vibration test system for spacecraft, specifically including:

[0062] A three-axis sinusoidal vibration test system for spacecraft includes: a main working platform, a connection assembly, multiple vibration tables, a cooling sub-module, an oil pump sub-module, a power amplifier sub-module, and a MIMO vibration control sub-module. Among them, the cooling sub-module is used to maintain the working temperatures of the main working platform and the power amplifier sub-module; the oil pump sub-module is used to provide a high-pressure oil source for the module and precisely control the vibration amplitude by adjusting the oil pressure; the power amplifier sub-module is used to amplify the low-power signal output by the MIMO control sub-module to the high-power level required by the main working platform; the MIMO vibration control sub-module is used to generate the excitation signals of each vibration table in real time based on the multi-input multi-output control algorithm.

[0063] The main working platform can refer to the core execution mechanism of vibration testing. It adopts a constant magnetic field and moving coil structure, generates an exciting force through an alternating current, and realizes the output of mechanical vibration energy. The magnitude of its exciting force is jointly determined by the current intensity, magnetic flux density, and effective length of the conductor. The connection assembly can refer to a mechanical structure for fixing the spacecraft under test and transmitting vibration energy. It usually includes a central holding device to ensure that the moving coil maintains a stable position under different loads and avoid dynamic drift. The vibration table can refer to a test device for simulating a mechanical vibration environment. It generates controllable vibration energy through electromechanical or hydraulic drive and is used to verify the structural strength, functional reliability, and fatigue life of products or materials under a vibration environment. In addition, each vibration table can be set at different positions to correspond to different test axes (such as the X / Y / Z axes in a preset three-dimensional coordinate system), and three-axis composite vibration can be achieved through synchronous control to simulate the complex working conditions of the spacecraft. The cooling sub-module can refer to a module for cooling the three-axis sinusoidal vibration test system for spacecraft. By circulating a cooling medium (such as water cooling or air cooling), it reduces the working temperatures of components such as the vibration table body and the power amplifier, preventing performance degradation or damage caused by overheating of the equipment. The oil pump sub-module can refer to a module that controls the amplitude accuracy of the vibration table by adjusting the flow rate and pressure of hydraulic oil and provides stable power support for the module at the same time. The power amplifier sub-module is usually an electric energy conversion device that can amplify the low-power signal output by the MIMO control sub-module to the high-power level required to drive the vibration table. It usually adopts linear power amplifier technology. The MIMO vibration control sub-module can refer to a multi-channel collaborative control core that generates the excitation signals of each vibration table in real time based on the multi-input multi-output algorithm and realizes closed-loop regulation through sensor feedback data to ensure precise control of the vibration magnitude and frequency.

[0064] A plurality of vibration tables are respectively arranged on one or both sides of the main working platform in the first, second or third direction, and are used to provide translational and / or rotational motions in the first, second or third direction to the main working platform, wherein the first direction is a preset direction in the horizontal plane, the second direction is a direction perpendicular to the first direction in the horizontal plane, and the third direction is a direction perpendicular to the horizontal plane.

[0065] The connecting assembly is arranged on the upper surface of the main working platform and is used to connect the main working platform and the spacecraft to be tested.

[0066] A plurality of three-axis force sensors are arranged on the connecting assembly. Among them, for each set three-axis force sensor, the set X-axis, set Y-axis and set Z-axis in the set three axes are respectively parallel to each other, and the set X-axis is parallel to the first direction, the set Y-axis is parallel to the second direction, and the set Z-axis is parallel to the third direction. A three-axis force sensor can refer to a high-precision sensing device that can simultaneously measure the forces or torques exerted on an object in three orthogonal directions (X, Y, Z axes), and is widely used in vibration testing, structural mechanics analysis and multi-axis dynamic load monitoring scenarios, providing key data support for the quantitative analysis of complex mechanical environments. As Figure 1 shown, the number of the three-axis force sensors can be 4 (the white squares 1, 2, 3, 4 in the figure). F x 、F y 、F z respectively represent the signals that can collect translational motions relative to the X-axis, Y-axis or Z-axis, and M x 、M y 、M z respectively represent the signals that can collect rotational motions of the X-axis, Y-axis or Z-axis. a can represent the vertical distance of the sensor from the center on the X-axis, and b can represent the vertical distance of the sensor from the center on the Y-axis.

[0067] The spacecraft is connected to the main working platform through the connecting assembly, and the spacecraft is located above the plurality of three-axis force sensors, and a full-scale vibration test is controlled for the plurality of vibration tables based on a preset requirement according to a preset three-axis sine test condition to test relevant parameters of the spacecraft.

[0068] In some optional implementation manners, the three-axis sine test condition includes:

[0069] The frequency value of each frequency point selected from M frequency points according to a set reference angle AN from a set frequency range [f s ,f p is determined according to a preset first calculation strategy, wherein the first calculation strategy is related to the number of octaves, and M = (360 / AN) 2-1. AN can be divisible by 90, and each frequency point is provided with a corresponding sequence number m of positive integers, where m ∈ [1, M]. The calculation of M can be further decomposed into M = [(90 / AN) * 4] 2 -1. As an example, when AN = 90°, M = (360 / 90) 2 -1 = 15. When AN = 45°, M = (360 / 45) 2 -1 = 63.

[0070] In some alternative implementation manners, the first calculation strategy includes:

[0071] Combined with a preset first calculation formula, calculate the frequency value of each frequency point, where the first calculation formula includes:

[0072]

[0073] where N represents the number of octaves, f m represents the frequency value of the m-th frequency point, represents the integer quotient of the dividend divided by the divisor, that is, the integer quotient after removing the remainder, and "dividend" mod "divisor" represents the integer remainder obtained by dividing the dividend by the divisor, that is, the integer remainder after removing the quotient.

[0074] The cyclic phase difference of the cyclic axis corresponding to each frequency point is calculated according to a preset second calculation strategy, and the distribution of the cyclic phase differences corresponding to all frequency points satisfies a sine distribution. Among them, the cyclic axis is one of the set X-axis direction, the set Y-axis direction, and the set Z-axis direction, and the cyclic phase difference is the phase difference of the cyclic axis corresponding to any frequency point with respect to the standard phase.

[0075] In some alternative implementation manners, the second calculation strategy includes:

[0076] Combined with a preset second calculation formula, calculate the cyclic phase difference of each frequency point, where the second calculation formula includes:

[0077]

[0078] where, represents the cyclic phase difference corresponding to the m-th frequency point.

[0079] The growth phase difference of the growth axis corresponding to each frequency point is calculated according to a preset third calculation strategy, where the growth axis is one of the set X-axis direction, the set Y-axis direction, and the set Z-axis direction that is not the cyclic axis.

[0080] In some alternative implementation manners, the third calculation strategy includes:

[0081] Calculate the growth phase difference at each frequency point in combination with a preset third calculation formula, where the third calculation formula includes:

[0082]

[0083] where represents the growth phase difference corresponding to the m-th frequency point.

[0084] The sum of the phase differences between any two of the three axes corresponding to each frequency point is 0, where the phase difference between axes is the phase difference between any two of the set three axes.

[0085] The design steps for the phase relationship conditions between axes are as follows:

[0086] Due to the mutual motion coupling effect between the vibrations in the three directions in three-axis vibration, the simple axial vector level conditions can no longer meet the needs of the test, and the mutual relationship between different axes also needs to be clarified. The three-axis sine test conditions are usually described by a matrix.

[0087] The designed three-axis sine test conditions must meet the requirements of physical space realizability, that is, the spectral matrix is required to be a positive definite matrix. Otherwise, it is impossible to find a set of suitable driving signals so that the response generated by exciting the structure meets the designed test condition reference spectral matrix. For a three-axis system, the response spectral matrix S can be expressed as follows:

[0088]

[0089] where E represents the identity matrix, X represents the data in the X-axis direction, Y represents the data in the Y-axis direction, Z represents the data in the Z-axis direction, X * represents the adjoint matrix of X, Y * represents the adjoint matrix of Y, Z * represents the adjoint matrix of Z, g * represents the adjoint matrix of g, represents the corresponding data between the *-axis and the *-axis.

[0090] Since the response cross-spectral matrix is a positive definite matrix, its determinant value must be greater than 0. Expand the determinant and simplify it to obtain:

[0091]

[0092] γ ** represents the coherence coefficient between the *-axis and the *-axis, represents the phase difference between the *-axis and the *-axis.

[0093] It mainly falls into the following three cases:

[0094] 1) Assume that the coherence coefficients between axes are all zero, that is Then the above inequality always holds, and the phase relationship between the axes has no constraints and can be any value. At this time, the reference spectral density matrix is a real diagonal matrix.

[0095] 2) Assume that the coherence coefficients between the axes are all 1, that is Then from the above inequality, it can be obtained that the sum of the phase differences between the three axes must be zero, that is

[0096]

[0097] 3) Assume that the sum of the phase differences between the axes is zero, then from the above inequality, it can be obtained that:

[0098]

[0099] The spectral density matrix can be used to quantify the motion of the product in space. The elements on the diagonal of the matrix contain the energy information of each control point, while the elements off the diagonal contain the information related to the correlation between each control point (such as the correlation coefficient and the phase difference). From the geometric constraint relationship, it can be known that: the closer the coherence coefficient between the control points is to 1, the more definite the relative motion between the control points; on the contrary, the closer the coherence coefficient between the control points is to 0, the more random the relative motion between the control points.

[0100] The beneficial effects of one of the embodiments in this specification at least include: conducting experiments on the above spacecraft three-axis sine vibration test system according to the preset three-axis sine test conditions, aiming to solve the problem that the spacecraft space motion trajectory is relatively single under the excitation of conventional three-axis sine test conditions, ensuring that the product is fully and widely excited from all angles in the three-dimensional space, making the spacecraft space motion fill the entire three-dimensional physical space, with a richer trajectory and more effective assessment. Moreover, compared with the traditional single-axis test method, it will greatly improve the mechanical test efficiency and mass production development efficiency of spacecraft products.

[0101] In some optional implementation manners, the three-axis sine test conditions further include: setting the amplitude magnitude conditions for each axis in the three axes as: based on interface force equivalence, and / or based on acceleration equivalence, and / or based on fatigue damage equivalence.

[0102] Specifically, the design method of three-axis six-degree-of-freedom vibration test conditions based on interface force equivalence (CN202210258221.3) is an invention patent. Starting from the resultant force and resultant moment of the spacecraft interface, it proposes a brand-new design method of three-axis six-degree-of-freedom test conditions based on interface force equivalence, and uses the three-axis sine test conditions to equivalently design the amplitude magnitude of each axis, aiming to solve the problem of designing the three-axis sine test conditions for spacecraft products, ensuring full assessment of the product and meeting the requirements of mechanical tests. Moreover, compared with the traditional single-axis test method, it greatly improves the mechanical test efficiency and mass production development efficiency of spacecraft products.

[0103] Multi-axis vibration test condition trimming method based on acceleration equivalence (CN201510797190.9) focuses on the acceleration response of key parts on the product. To ensure that the response in each direction during multi-axis vibration is the same as that when uniaxial vibrations are applied separately, a multi-axis vibration test amplitude level condition trimming method based on acceleration equivalence is designed with the uniaxial vibration test conditions as the benchmark.

[0104] Multi-axis vibration test condition trimming method based on fatigue damage equivalence analysis (CN201611047327.X), mainly starting from fatigue damage equivalence, can ensure that the actual stress fatigue situation can be truly simulated during multi-axis vibration tests, thus simulating the real test environment. At the same time, the positions on the product that are difficult to assess in the actual environment can be assessed through this method.

[0105] The design methods for the amplitude level conditions in each axial direction can refer to the above three patents and be designed in detail from three aspects: interface force equivalence, acceleration equivalence, and fatigue damage equivalence, and will not be elaborated here.

[0106] In some optional implementation manners, multiple three-axis force sensors are connected to a preset signal conditioner. The signal conditioner is provided with multiple signal channels. Among them, p signal channels of the signal conditioner are respectively connected to the set Z axes of each three-axis force sensor, where the number of three-axis force sensors is p. Another p signal channels of the signal conditioner are respectively connected to the combined signals of the set X axes or Y axes of any two three-axis force sensors.

[0107] In some optional implementation manners, the number of vibration tables is 8. Among them, determining the first axis, the second axis, and the third axis that pass through the geometric centroid of the main working platform and are respectively parallel to the first direction, the second direction, and the third direction specifically includes:

[0108] One vibration table with a vibration direction parallel to the first axis is arranged on each side of the first axis to provide a vibration source parallel to the first axis.

[0109] One vibration table with a vibration direction parallel to the second axis is arranged on each side of the second axis to provide a vibration source parallel to the second axis.

[0110] Four vibration tables arranged in a pairwise symmetric rectangular array around the third axis and with vibration directions parallel to the third axis are provided to provide a vibration source parallel to the third axis. Setting more vibration tables parallel to the third axis can simulate more vibrations in the third direction (i.e., the Z-axis direction) to perform more targeted detection on spacecraft.

[0111] In some alternative implementations, multiple measurement points are provided on the surface and / or inside of the spacecraft. The measurement points are used to collect vibration response data of various parts of the spacecraft, evaluate the structural dynamics characteristics, and check for abnormal responses to assist in fault diagnosis.

[0112] To further explain the embodiments of the present invention, the following provides a detailed embodiment to further explain the present disclosure.

[0113] In this embodiment, the reference angle AN is set to 90°, the frequency range is [f s , f p , the number of octaves is N (refer to the aforementioned calculation method), the strategy is related to the number of octaves, M = (360 / AN) 2 - 1 = 15, each frequency point is provided with a corresponding positive integer sequence number m, m ∈ [1, 15], corresponding to f2 - f in the following table 16 . The phase differences between the three axes corresponding to each frequency point are respectively and Let the Z-axis be the cycle axis and the Y-axis be the growth axis. The parameters of each frequency point are shown in Table 1 below:

[0114] Table 1 Introduction to the relevant parameters of each frequency point

[0115]

[0116]

[0117] Figure 2 , Figure 3 , Figure 4 are respectively another schematic diagram of the test conditions for the phase parameters of the X / Y / Z three-axis sinusoidal vibration of the spacecraft.

[0118] In addition, the output cables of the three-axis force sensors must be correctly connected to the signal conditioner. The wiring methods between the three-axis force sensor cables and the signal conditioner are listed in the second column of Table 2. For the first four channels of the signal conditioner, the signals connected to each channel are the combined force signals from two different three-axis force sensors; for the last four channels, the signals connected to each channel are respectively the signals of a single three-axis force sensor in the set Z-axis direction.

[0119] Table 2 Output of the force signal conditioner for the four force sensors

[0120]

[0121]

[0122] Wherein, Ch** represents the serial number of the **th signal channel, FX represents the signal in the X-axis direction of the connection, FY represents the signal in the Y-axis direction of the connection, and FZ represents the signal in the Z-axis direction of the connection. a, b, and K respectively represent preset calculation coefficients.

[0123] In some embodiments, the embodiment of the present invention further provides a method for the three-axis sine vibration test of a spacecraft. The method is applied to the aforementioned three-axis sine vibration test system for a spacecraft, and the method includes:

[0124] Sequentially turn on the main vibration table, multiple vibration tables, cooling sub-module, oil pump sub-module, power amplifier sub-module, and MIMO vibration control sub-module of the module, check whether the operating states of all components are normal, and set the main vibration table at the coarse and medium position.

[0125] Use a sling to lift the spacecraft to be tested, position the spacecraft above the connection assembly, slowly lower it, and then fixedly connect the spacecraft to the connection assembly. Remove the sling. Among them, the spacecraft to be tested is initially set to the power-off state.

[0126] Conduct precise centering debugging on the main vibration table.

[0127] Turn on multiple three-axis force sensors, perform a pre-vibration level conduction vibration test based on a preset first parameter, perform data processing and recording through a preset data processing module after self-checking, and detect the corresponding data of each measurement point and perform anomaly processing after the pre-vibration level conduction vibration test ends.

[0128] Set the spacecraft to be tested to the power-on state, turn on multiple three-axis force sensors, perform the first characteristic level vibration test based on a preset second parameter, perform data processing and recording through the data processing module after self-checking, monitor the working state of the spacecraft to be tested throughout the test process, and detect the corresponding data of each measurement point and perform anomaly processing after the pre-vibration level conduction vibration test ends.

[0129] Turn on multiple three-axis force sensors, perform a full-scale vibration test based on preset three-axis sine test conditions, perform data processing and recording through the data processing module after self-checking, monitor the working state of the spacecraft to be tested throughout the test process, and evaluate the effectiveness of the three-axis sine test conditions after the full-scale vibration test ends.

[0130] Set the spacecraft under test to the power-on state, turn on multiple three-axis force sensors, perform the second characteristic-level vibration test based on a preset third parameter, and after self-check, process and record the data through a data processing module. Monitor the working state of the spacecraft under test throughout the test process, compare the relevant data in the two characteristic-level vibration tests, determine whether the test results of the two characteristic-level vibration tests are consistent, and check whether there is any damage or other abnormality in the appearance of the spacecraft under test to determine whether the spacecraft under test can safely withstand the sine vibration test.

[0131] Dismantle the spacecraft under test and lift it off, and sequentially turn off the main vibration table, multiple vibration tables, cooling sub-module, oil pump sub-module, power amplifier sub-module, and MIMO vibration control sub-module, and the test is over.

[0132] In some embodiments, the specific implementation and the technical effects brought by the foregoing steps corresponding to those in the corresponding embodiments of the foregoing spacecraft three-axis sine vibration test system can refer to those steps corresponding to the foregoing spacecraft three-axis sine vibration test system, and will not be elaborated here.

[0133] In the above embodiments, the descriptions of each embodiment have their own emphases. For the parts not detailed in a certain embodiment, reference can be made to the relevant descriptions of other embodiments.

[0134] The preferred embodiments of the present specification disclosed above are only used to help explain the present specification. The alternative embodiments do not describe all the details in detail, nor do they limit the invention to only the specific implementation manners. Obviously, many modifications and changes can be made according to the content of the embodiments of the present specification. The present specification selects and specifically describes these embodiments to better explain the principles and practical applications of the embodiments of the present specification, so that those skilled in the art can understand and utilize the present specification well. The present specification is only limited by the claims and their full scope and equivalents.

Claims

1. A three-axis sine vibration test system for a spacecraft, characterized in that, Including: A main working platform, a connection component, multiple vibration tables, a cooling sub-module, an oil pump sub-module, a power amplifier sub-module, and a MIMO vibration control sub-module, where the cooling sub-module is used to maintain the working temperatures of the main working platform and the power amplifier sub-module; the oil pump sub-module is used to provide a high-pressure oil source for the module and precisely control the vibration amplitude by adjusting the oil pressure; the power amplifier sub-module is used to amplify the low-power signal output by the MIMO control sub-module to the high-power level required by the main working platform; the MIMO vibration control sub-module is used to generate excitation signals for each vibration table in real time based on the multi-input multi-output control algorithm; the multiple vibration tables are respectively arranged on one side or both sides of the main working platform in the first, second, or third direction, and are used to provide translational and / or rotational motions in the first, second, or third direction for the main working platform, where the first direction is a preset direction in the horizontal plane, the second direction is a direction perpendicular to the first direction in the horizontal plane, and the third direction is a direction perpendicular to the horizontal plane; the connection component is arranged on the upper surface of the main working platform and is used to connect the main working platform and the spacecraft to be tested; a plurality of three-axis force sensors are arranged on the connection component, and for each set three-axis force sensor, the set X-axis, set Y-axis, and set Z-axis in the set three axes are respectively parallel to each other, and the set X-axis is parallel to the first direction, the set Y-axis is parallel to the second direction, and the set Z-axis is parallel to the third direction; the spacecraft is connected to the main working platform through the connection component, the spacecraft is located above the plurality of three-axis force sensors, and controls the plurality of vibration tables to perform a full-scale vibration test based on preset requirements according to preset three-axis sine test conditions to test relevant parameters of the spacecraft.

2. The system according to claim 1, wherein The three-axis sine test conditions include: According to the set reference angle AN from the set frequency range [f s , f p , the frequency value of each of the M frequency points screened is determined according to a preset first calculation strategy, where the first calculation strategy is related to the number of octaves, M = (360 / AN) 2 -1, and each frequency point is provided with a corresponding sequence number m of positive integers, m ∈ [1, M]; the cyclic phase difference of the cyclic axis corresponding to each frequency point is calculated according to a preset second calculation strategy, and the distribution of the cyclic phase differences corresponding to all frequency points satisfies a sine distribution, where the cyclic axis is one of the set X-axis, set Y-axis, and set Z-axis, and the cyclic phase difference is the phase difference of the cyclic axis corresponding to any frequency point with respect to the standard phase; the growth phase difference of the growth axis corresponding to each frequency point is calculated according to a preset third calculation strategy, where the growth axis is one of the non-cyclic axes among the set X-axis, set Y-axis, and set Z-axis; the sum of the inter-axis phase differences of the three axes corresponding to each frequency point is 0, where the inter-axis phase difference is the phase difference between any two of the set three axes.

3. The system according to claim 2, wherein The first calculation strategy includes: combining a preset first calculation formula to calculate the frequency value of each frequency point, where the first calculation formula includes: where N represents the number of octaves, and f m represents the frequency value of the m-th frequency point, represents the integer quotient of the dividend divided by the divisor, and "dividend" mod "divisor" represents the integer remainder obtained by dividing the dividend by the divisor.

4. The system according to claim 2, wherein The second calculation strategy includes: combining a preset second calculation formula to calculate the cyclic phase difference of each frequency point, where the second calculation formula includes: Among them, represents the cyclic phase difference corresponding to the m-th frequency point.

5. The system according to claim 2, wherein The third calculation strategy includes: combining a preset third calculation formula to calculate the growth phase difference of each frequency point, where the third calculation formula includes: Among them, represents the growth phase difference corresponding to the m-th frequency point.

6. The system according to claim 1, wherein The three-axis sine test conditions further include: The amplitude magnitude conditions for each axis in the set three axes are: Based on the equivalence of interface forces, and / or Based on the equivalence of acceleration, and / or Based on the equivalence of fatigue damage.

7. The system according to any one of claims 1 to 6, characterized in that The multiple three-axis force sensors are connected to a preset signal conditioner, and the signal conditioner is provided with multiple signal channels. Among them, p signal channels of the signal conditioner are respectively connected to the set Z axes of each of the three-axis force sensors, where the number of the three-axis force sensors is p; Another p signal channels of the signal conditioner are respectively combined with the set X axes or Y axes of any two of the three-axis force sensors.

8. The system according to any one of claims 1 to 6, characterized in that, The number of the vibration tables is 8, where Determine the first axis, the second axis, and the third axis that pass through the geometric centroid of the main working platform and are respectively parallel to the first direction, the second direction, and the third direction; One vibration table with a vibration direction parallel to the first axis is arranged on each side of the first axis; One vibration table with a vibration direction parallel to the second axis is arranged on each side of the second axis; Four vibration tables are arranged circumferentially around the third axis in a pairwise symmetric rectangular arrangement, and the vibration direction is parallel to the third axis.

9. The system according to any one of claims 1 to 6, characterized in that, Multiple measuring points are arranged on the surface and / or inside of the spacecraft. The measuring points are used to collect the vibration response data of each part of the spacecraft, evaluate the structural dynamic characteristics, and check for abnormal responses to assist in fault diagnosis.

10. A three-axis sinusoidal vibration test method for a spacecraft, characterized in that, The method is applied to the spacecraft three-axis sine vibration test system according to any one of claims 1 to 9. The method includes: Sequentially turn on the main vibration table, multiple vibration tables, the cooling sub-module, the oil pump sub-module, the power amplifier sub-module, and the MIMO vibration control sub-module of the system, check whether the operating states of all components are normal, and set the main vibration table to the rough and medium position; Use a sling to lift the spacecraft to be tested, place the spacecraft above the connection assembly, slowly lower it, and then fixedly connect the spacecraft to the connection assembly, and remove the sling. Among them, the spacecraft to be tested is initially set to the power-off state; Perform precise centering debugging on the main vibration table; Turn on multiple three-axis force sensors, perform a pre-vibration level conduction vibration test based on a preset first parameter, perform self-check, and then process and record the data through a preset data processing module. After the pre-vibration level conduction vibration test is completed, detect the corresponding data of each measuring point and perform abnormal processing; Set the spacecraft to be tested to the power-on state, turn on multiple three-axis force sensors, perform the first characteristic level vibration test based on a preset second parameter, perform self-check, and then process and record the data through the data processing module. Monitor the working state of the spacecraft to be tested throughout the test process, and after the pre-vibration level conduction vibration test is completed, detect the corresponding data of each measuring point and perform abnormal processing; Turn on multiple three - axis force sensors, perform a full - scale vibration test based on the preset three - axis sine test conditions, process and record the data through the data processing module after self - inspection, monitor the working state of the spacecraft under test throughout the test process, and evaluate the effectiveness of the three - axis sine test conditions after the full - scale vibration test is completed; Set the spacecraft under test to the power - on state, turn on multiple three - axis force sensors, perform a second characteristic - level vibration test based on the preset third parameter, process and record the data through the data processing module after self - inspection, monitor the working state of the spacecraft under test throughout the test process, compare the relevant data in the two characteristic - level vibration tests, judge whether the test results of the two characteristic - level vibration tests are consistent, and check whether there is any damage or other abnormality in the appearance of the spacecraft under test to determine whether the spacecraft under test can safely withstand the sine vibration test; Dismantle the spacecraft under test and lift it off, turn off the main vibration table, multiple vibration tables, cooling sub - module, oil pump sub - module, power amplifier sub - module and MIMO vibration control sub - module in sequence, and the test is over.

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