CN114492142A - Device and method for testing fire impact resistance of spacecraft component - Google Patents

Abstract

The invention discloses a device and a method for testing the fire impact resistance of spacecraft components. The test device comprises a test bracket, a bearing piece, an impact source and an acquisition system; the bearing part is freely suspended on the test bracket and is used for bearing the spacecraft component to be tested; the impact source is used for providing laser to generate laser impact on one surface, facing the laser, of the bearing piece, and the impact is transmitted to the installation position of the spacecraft component to be tested, which is arranged on the other surface of the bearing piece, through the bearing piece, so that the simulation of the real fire impact environment of the spacecraft component is realized; the acquisition system is connected with the bearing piece and is used for acquiring the physical parameters of vibration response of a point of the attachment of the installation position where the spacecraft component is located after the point is impacted by laser; the invention can well represent the high-frequency fire impact characteristic, and realize the accurate simulation of the real fire impact environment of the spacecraft component, so as to test the fire impact resistance of the spacecraft component.

Description

Apparatus and method for testing the pyrotechnic shock resistance of spacecraft components

technical field

The invention relates to the technical field of pyrotechnic impact environment simulation and pyrotechnic impact resistance testing of spacecraft components, and mainly relates to a device and method for testing the pyrotechnic impact resistance capabilities of spacecraft components.

Background technique

The aerospace pyrotechnic shock environment is the structural shock response caused by the explosive separation of pyrotechnics in the process of star-rocket separation, fairing separation, inter-stage separation, and component deployment. The aerospace pyrotechnic shock environment has the characteristics of high frequency, transient state and high magnitude, which can easily cause damage to the precise electronic components and other instruments and equipment on the spacecraft. The aerospace pyrotechnic shock environment is one of the most severe mechanical environments experienced by the spacecraft during the entire life cycle. During the development of the spacecraft, it is necessary to conduct a pyrotechnic shock simulation test on the ground in advance, that is, the fire resistance of the spacecraft components. Work impact ability is assessed and tested. Up to now, the simulation experiment methods for the resistance to pyrotechnic shock are mainly divided into two types: pyrotechnic explosion type and non-pyrotechnic explosion type.

The former directly uses gunpowder or detonator, which is loaded on different types of resonance devices in a certain way, and obtains the pyrotechnic impact response by detonating the gunpowder or detonator. This simulation method is a destructive excitation test, which has poor operability and repeatability of results. Before the formal test, many trials and errors are often required to obtain shock response test results with relatively high relative accuracy. The test cost is high, and the safety The performance is poor, and it is easy to cause great damage to the test personnel and the test equipment. The latter mainly includes two types of mechanical vibration type and vibrating table. The pendulum, air gun or vibration table itself is used as the impact generator, and its vibration response is obtained through the resonance device such as the resonance plate and the resonance rod to simulate the spacecraft components. Pyrotechnical shock response under real conditions.

The non-fire explosive test method has good operability and repeatability, and the test cost is low, but its frequency spectrum is narrow (the frequency spectrum of the vibration table can only reach about 3000Hz), the high frequency response amplitude is small, and it cannot simulate fire The high frequency response characteristics of the shock (its main frequency is between 100Hz-100Khz).

SUMMARY OF THE INVENTION

The present invention aims to solve at least one of the technical problems existing in the prior art. Therefore, an object of the present invention is to provide a device for testing the pyrotechnic impact resistance of spacecraft components, which can simulate the spacecraft pyrotechnic impact environment and test the pyrotechnic impact resistance of spacecraft components.

The device for testing the pyrotechnic impact resistance of spacecraft components according to the embodiment of the first aspect of the present invention includes:

test stand;

a carrier, which is freely suspended on the test bracket and used to carry the spacecraft components to be tested;

an impact source, the impact source is used for providing laser light to generate laser shock on the side of the carrier facing the laser, the laser shock is transmitted to the aerospace aircraft disposed on the other surface of the carrier via the carrier At the installation position of the spacecraft components, the simulation of the real pyrotechnic impact environment of the spacecraft components is realized;

A collection system, which is connected to the carrier, is used to collect the physical parameters of the vibration response of an attachment point at the installation position of the spacecraft component after being impacted by the laser.

The device for testing the pyrotechnic impact resistance of spacecraft components according to the embodiment of the first aspect of the present invention has a simple structure and simple operation method, can well characterize the high-frequency pyrotechnic impact characteristics, and realize the real fire impact of spacecraft components. Simulation of the industrial shock environment in order to test the thermal shock resistance of spacecraft components.

In some embodiments, an optical path control system is further included, the optical path control system is arranged between the impact source and the carrier, and is used for controlling the optical path of the laser light transmission.

In some embodiments, the optical path control system includes a reflecting mirror and a focusing lens, and after the laser light emitted by the impact source is reflected by the reflecting mirror to the focusing lens, the focusing lens focuses and irradiates the laser light on the focusing lens. on the carrier.

In some embodiments, the focal length of the focusing lens is set to be adjustable.

In some embodiments, the carrier includes a base layer, an absorbing coating disposed on a side of the base layer facing the laser, and a transparent confinement layer disposed on the absorbing coating;

When the laser passes through the transparent confinement layer and reaches the absorption coating, the absorption coating absorbs the energy of the laser and generates plasma, and the plasma absorbs the laser under the confinement of the transparent confinement layer. The remaining energy of the laser expands rapidly (that is, the phenomenon of plasma explosion), and the shock wave is transmitted through the carrier to the installation position of the spacecraft component to be tested fixed on the other surface of the carrier, so as to realize the detection of the spacecraft. Simulation of component pyrotechnic shock environments.

In some embodiments, a flexible member is further included, one end of the flexible member is fixed to the test bracket, and the other end of the flexible member is fixed to the bearing member, so that the bearing member is freely suspended on the test bracket.

In some embodiments, the incident angle of the laser light relative to the carrier is set to an adjustable predetermined angle.

In some embodiments, the acquisition system includes an acceleration sensor, and the acceleration sensor is attached near the installation position of the carrier spacecraft components, but should not interfere with the spacecraft components.

The second aspect of the present invention also provides a method for testing the pyrotechnic impact resistance of spacecraft components.

According to the method for testing the pyrotechnic shock resistance of a spacecraft component according to an embodiment of the second aspect of the present invention, the method for testing the pyrotechnic impact resistance of a spacecraft component according to any one of the embodiments of the first aspect of the present invention is adopted. The device includes the following steps:

S1: Install the spacecraft components to be tested at the specified installation position on the carrier by pasting;

S2: start the shock source, and perform one or more laser shocks on the carrier according to actual needs;

S3: utilize the acquisition system to collect the physical parameters of the vibration response after being subjected to the laser shock effect at a point near the installation position of the spacecraft components, and determine the laser shock intensity suffered by the spacecraft components to be tested;

S4: If the laser shock strength does not meet the actual pyrotechnic shock environment test requirements, adjust the output energy of the shock source and the installation position of the spacecraft components; then repeat steps S1 to S4 to make the laser shock strength finally meet the requirements of the spacecraft components The actual pyrotechnic shock environment test requirements of the device;

S5: Inspect the normal use performance of the spacecraft components after the laser shock, so as to test the pyrotechnic shock resistance of the spacecraft components to be tested.

To sum up, steps S1 to S4 are the simulation process of the actual pyrotechnic impact environment of the components of the spacecraft to be tested; Test of components' resistance to pyrotechnic shock.

According to the method for testing the pyrotechnic shock resistance of spacecraft components according to the embodiment of the second aspect of the present invention, the operation method is simple and convenient. Test of pyrotechnic impact capability.

In some embodiments, the physical parameter includes an acceleration time domain response signal of a point near the installation position of the spacecraft component after being subjected to the laser shock.

In some embodiments, the following steps are also included:

analyzing the acceleration time-domain response signal, and drawing an impact response spectrum;

Taking the response time-domain peak value of the acceleration time-domain response signal as the shock-response spectrum amplitude of the shock-response spectrum at a predetermined frequency.

In some embodiments, the following step is further included: using the same shock source to test the repeatability and controllability of the test results.

In some embodiments, collecting laser shock acceleration time-domain response signals for multiple tests;

Decomposing the laser shock acceleration time-domain response signal of the multiple tests into a single laser shock acceleration time-domain response signal according to a time series;

drawing the single-shot laser shock acceleration time-domain response signal into the shock response spectrum;

converting the shock response spectrum into an image representing shock frequencies in pixel values;

Convert the image to an impulse response spectrogram.

In some embodiments, the method further includes the following steps: establishing a laser shock finite element model of a spacecraft component; and performing a dynamic numerical simulation on the impact response of the spacecraft component.

Additional aspects and advantages of the present invention will be set forth, in part, from the following description, and in part will be apparent from the following description, or may be learned by practice of the invention.

Description of drawings

The above and/or additional aspects and advantages of the present invention will become apparent and readily understood from the following description of embodiments taken in conjunction with the accompanying drawings, wherein:

FIG. 1 is a schematic diagram of an apparatus for testing the pyrotechnic shock resistance of a spacecraft component according to an embodiment of the first aspect of the present invention.

FIG. 2 is a partial enlarged view of the carrier area of FIG. 1 .

FIG. 3 is a schematic diagram of another apparatus for testing the pyrotechnic shock resistance of spacecraft components according to an embodiment of the first aspect of the present invention.

FIG. 4 is a partial enlarged view of the carrier area of FIG. 3 .

FIG. 5 is a schematic diagram of a method for testing the pyrotechnic shock resistance of a spacecraft component according to an embodiment of the second aspect of the present invention.

6a to 6c are schematic diagrams of different included angles of the rotation of the carrier in the method for testing the thermal shock resistance of a spacecraft component according to an embodiment of the second aspect of the present invention.

FIG. 7 is a method for testing the pyrotechnic shock resistance of a spacecraft component according to an embodiment of the second aspect of the present invention. The apparatus for testing the pyrotechnic impact resistance of a spacecraft component according to the first aspect of the present invention is shown in FIG. Schematic diagram of the shock response spectrum of the pyrotechnic shock environment obtained under laser shock.

FIG. 8 is a schematic diagram of the dynamic model of the single-degree-of-freedom mass-spring-damper system required to synthesize the shock response spectrum shown in FIG. 6 .

FIG. 9 is an impulse response spectrum measured by the device shown in FIG. 1 .

Fig. 10 shows the relationship between the laser energy density measured by the device shown in Fig. 1 and the relative coefficient of the impulse response spectrum amplitude.

FIG. 11 is a schematic front view of the finite element model in the method for testing the pyrotechnic shock resistance of a spacecraft component according to an embodiment of the second aspect of the present invention.

12 is a schematic side view of a finite element model in the method for testing the pyrotechnic impact resistance of a spacecraft component according to an embodiment of the second aspect of the present invention.

FIG. 13 is a partial enlarged view of the load-carrying region of the bearing member region of FIG. 11 .

Figure 14 is a schematic diagram of the pressure load at the load-loading part of the finite element model of the bearing member.

Reference number:

Apparatus 1000 for testing the pyrotechnic shock resistance of spacecraft components

Test stand 1

carrier 2 base layer 201 absorbing coating 202 transparent constraining layer 203

Shock source 3 Laser 301

Acquisition System 4 Accelerometer 401

Optical path control system 5 Reflector 501 Focusing lens 502

Spacecraft Components 6 Flexible Parts 7 Workbench 8

Carrier Finite Element Model 901

Spacecraft Components Finite Element Model 902

Bearing part finite element model load loading position 903

Detailed ways

The following describes in detail the embodiments of the present invention, examples of which are illustrated in the accompanying drawings, wherein the same or similar reference numerals refer to the same or similar elements or elements having the same or similar functions throughout. The embodiments described below with reference to the accompanying drawings are exemplary, only used to explain the present invention, and should not be construed as a limitation of the present invention.

The following describes an apparatus 1000 and a method for testing the pyrotechnic shock resistance of a spacecraft component according to an embodiment of the present invention with reference to FIGS. 1 to 14 .

As shown in FIGS. 1 to 4 , the apparatus 1000 for testing the pyrotechnic shock resistance of spacecraft components according to the embodiment of the first aspect of the present invention can simulate the pyrotechnic impact environment of the spacecraft, and test the spacecraft components 6's resistance to fire shock.

The apparatus 1000 for testing the pyrotechnic impact resistance of spacecraft components according to the embodiment of the first aspect of the present invention includes a test bracket 1 , a carrier 2 , an impact source 3 and a collection system 4 . Among them, the carrier 2 is freely suspended on the test bracket 1, and the carrier 2 is used to carry the spacecraft components 6 to be tested; the shock source 3 is used to provide the laser 301 to generate laser shock on the side of the carrier 2 facing the laser. The laser shock is transmitted through the carrier 2 to the installation position of the spacecraft component 6 set on the other surface of the carrier 2 to simulate the real pyrotechnic impact environment of the spacecraft component 6; the acquisition system 4 and the carrier 2 is connected, and is used to collect the physical parameters of the vibration response of a point near the installation position of the spacecraft component 6 after being impacted by the laser 301 .

Specifically, the test stand 1 is a stand structure, so as to freely hang and fix the carrier 2 so that the carrier 2 is in a suspended state (eg, as shown in FIG. 1 and FIG. 2 ), so as to facilitate the experiment.

The carrier 2 is freely suspended on the test bracket 1, and the carrier 2 is used to carry the spacecraft components 6 to be tested; that is, the carrier 2 is fixed on the test stand in a suspended state by means of free suspension on the one hand. On the other hand, the bracket 1 is used for fixing with the spacecraft component 6 to be tested, so as to realize the simulation of the free boundary of the pyrotechnic impact of the spacecraft component.

The shock source 3 is used to provide the laser 301 to generate laser shock on the side of the carrier 2 facing the laser, and the laser shock is transmitted through the carrier 2 to the installation position of the spacecraft component 6 fixed on the other surface of the carrier 2, Realize the simulation of the real pyrotechnic impact environment of the spacecraft components 6. Specifically, the impact source 3 can be a laser generator, and the structure of the laser generator includes a power supply system, a laser generator, a cooling system, etc., preferably the laser generator is a solid-state laser generator, which can generate high energy density, narrow When the laser 301 provided by the shock source 3 acts on a surface of the carrier 2 facing the laser, the surface of the carrier 2 (that is, the surface facing the laser) will have a plasma explosion phenomenon, and produce The shock wave is transmitted through the carrier 2 to the installation position of the spacecraft component 6 fixed on the other surface of the carrier 2 (the surface is the surface facing away from the laser), so that the spacecraft component 6 is generated under the conditions of real pyrotechnic shock. Very similar high-frequency vibration response, so as to realize the simulation of the real pyrotechnic shock environment where the spacecraft components 6 are located.

The acquisition system 4 is connected to the carrier 2 and is used to acquire the physical parameters of the vibration response at a point near the installation position of the spacecraft component 6 after being impacted by the laser 301 . Since equipment is already installed at the installation position of the spacecraft component 6, it is impossible to directly measure the physical parameters. In the actual measurement process, a point near the installation position of the spacecraft component 6 is usually used for approximate substitution. Through the analysis and processing of the physical parameters, the intensity of the pyrotechnic shock environment where the spacecraft components 6 are located can be obtained, and the normal use performance of the spacecraft components 6 after laser shock can be tested to realize the detection of the spacecraft components. Test of the resistance to pyrotechnic shock of device 6.

When the spacecraft component 6 is tested by using the apparatus 1000 for testing the pyrotechnic impact resistance of the spacecraft component according to the embodiment of the first aspect of the present invention, the impact source 3 can be placed on the workbench 8, and the impact source 3 A test bracket 1 is provided on one side; the spacecraft component 6 to be tested is fixed at the specified installation position on the carrier 2; the shock source 3 is activated to perform one or more laser shocks on the carrier 2; the acquisition system 4 is used The physical parameters of the vibration response of a point near the installation position of the spacecraft component 6 after being subjected to laser shock are collected; the normal use performance of the spacecraft component 6 to be tested after the laser shock is tested.

The device 1000 for testing the pyrotechnic impact resistance of spacecraft components according to the embodiment of the first aspect of the present invention has a simple structure and simple operation method, can well characterize the high-frequency pyrotechnic impact characteristics, and realizes the detection of spacecraft components. The simulation of the real pyrotechnic shock environment of the device is used to test the pyrotechnic shock resistance of the spacecraft components.

In some embodiments, as shown in FIG. 1 and FIG. 3 , an optical path control system 5 is further included, and the optical path control system 5 is arranged between the impact source 3 and the carrier 2 for controlling the optical path of the laser light transmission. That is to say, by setting the optical path control system 5, after the laser 301 emitted by the shock source 3 is focused, the energy of the laser 301 is more concentrated, which can effectively ensure that the focused laser 301 can generate the required amount of the experiment when it acts on the carrier 2. level shock wave. It should be noted that the optical path control system 5 is generally placed on the workbench 8, but is not limited by this, and can be adjusted according to the actual situation.

In some embodiments, as shown in FIG. 3 , the optical path control system 5 includes a reflecting mirror 501 and a focusing lens 502 . After the laser light 301 emitted by the impact source 3 is reflected by the reflecting mirror 501 to the focusing lens 502 , the focusing lens 502 focuses and irradiates on the focusing lens 502 . on the carrier 2. The function of the mirror 501 is to change the propagation path of the laser light 301 emitted by the shock source 3 , and the function of the focusing lens 502 is to adjust the spot diameter of the laser light 301 emitted by the shock source 3 .

In some embodiments, the focal length of focusing lens 502 is set to be adjustable. In this way, the focal length of the focusing lens 502 can be set according to the requirements of the experiment to adjust the size of the light spot.

In some embodiments, as shown in FIGS. 2 and 4 , the carrier 2 includes a base layer 201 , an absorbing coating 202 disposed on the side of the base layer 201 facing the laser 301 , and a transparent constraint disposed on the absorbing coating 202 Layer 203; when the laser 301 passes through the transparent confinement layer 203 and reaches the absorption coating 202, the absorption coating 202 absorbs the energy of the laser 301 and generates plasma, and the plasma absorbs the remaining energy of the laser 301 under the confinement of the transparent confinement layer 203 The rapid expansion (ie, the plasma explosion phenomenon) generates shock waves and is transmitted through the carrier 2 to the spacecraft component 6 fixed on the other surface of the carrier 2 .

Specifically, an absorption coating 202 and a transparent confinement layer 203 are provided on the side of the base layer 201 facing the laser 301 , wherein the absorption coating 202 is located between the base layer 201 and the transparent confinement layer 203 , and the other side of the base layer 201 The spacecraft component 6 to be tested can be fixed. The provision of the transparent confinement layer 203 and the absorption coating 202 can effectively enhance the impact pressure of the laser 301 and prolong the duration of the laser 301 impact. After the laser 301 emitted by the shock source 3 is focused by the optical path control system 5 , it first passes through the transparent confinement layer 203 . Under the ionization action of the high-energy-density laser 301, the absorption coating 202 will absorb the energy of the laser 301 and generate plasma. Under the constraint of the transparent confinement layer 203, the plasma absorbs the remaining energy of the laser 301 and rapidly expands (that is, the plasma). explosion phenomenon), a shock wave is generated and transmitted through the carrier 2 to the spacecraft component 6 fixed on the other surface of the carrier 2 . Among them, the installation position of the spacecraft component 6 is usually located on the side of the carrier 2 facing away from the laser 301, and close to the incident position of the laser 301, so as to improve the magnitude of the laser impact, but it is not limited to this, and can be adjusted according to actual needs.

Optionally, the base layer 201 may be made of materials such as aluminum alloy, titanium alloy, or stainless steel, but it is not limited to this, and can be adjusted according to actual needs.

The absorption coating 202 can be aluminum foil or black tape, and its position is set on the side of the base layer 201 facing the laser 301 . Meanwhile, the thickness of the absorption coating 202 is set to be very thin, and the thickness may be 0.1 mm, but it is not limited to this, and can be adjusted according to actual needs.

The transparent constraining layer 203 may be distilled water or transparent glass. Among them, the impact effect of distilled water to enhance the laser 301 is worse than that of transparent glass, but distilled water as the transparent constraining layer 203 has the advantages of convenient use and good repeatability of experimental results. Therefore, the material of the transparent constraining layer 203 is preferably distilled water, and the schematic diagram of the structure is shown in FIG. 1 . In addition, when the transparent constraining layer 203 is made of distilled water, the testing device further includes a water spray pipe connected to an external water source. The distilled water is sprayed from the water spray pipe to form a transparent constraining layer 203 with a certain thickness on the carrier 2 . In addition, it is worth noting that when the transparent confinement layer 203 is distilled water, part of the distilled water near the absorption coating 202 will also be ionized under the action of the laser 301 .

In some embodiments, as shown in FIG. 1 , a flexible member 7 is further included. One end of the flexible member 7 is fixed on the test bracket 1 , and the other end is fixed on the carrier 2 , so that the carrier 2 is freely suspended from the test bracket 1 . Therefore, the flexible member 7 is used to fix the bearing member 2 on the test bracket 1, the assembly operation is very simple, and the position of the bearing member can be adjusted. For example, as shown in FIG. 1 , the carrier 2 can be freely suspended from the test stand 1 through four flexible members 7 , and the height of the carrier 2 can be adjusted by changing the length of the flexible members 7 . In addition, it is worth noting that the number of flexible members 7 used is not limited to four, and can be selected according to actual needs.

In some embodiments, the incident angle of the laser light 301 relative to the carrier 2 is set to an adjustable predetermined angle. At this time, the incident angle can be understood as the angle between the laser light 301 and the carrier 2 . Changing the angle between the laser 301 and the carrier 2 will affect the shape of the spot of the laser 301 and the energy distribution of the laser 301 , thereby affecting the shock wave impact pressure of the laser 301 . Therefore, the present invention also sets the included angle between the laser 301 and the carrier 2 to different predetermined angles, and collects the laser 301 received at the installation position of the spacecraft component to be tested under different predetermined angles (as shown in FIG. 6a to FIG. 6c ). The physical parameters of the vibration response after the impact, quantitatively determine the influence law of the angle between the laser 301 and the carrier 2 on the pyrotechnic impact environment at the installation location of the spacecraft components 6, so as to accurately adjust the spacecraft components to be tested 6 The pyrotechnic shock environment. Wherein, by rotating the angle of the carrier 2 (as shown in FIGS. 6 a to 6 c ), the angle between the laser 301 and the carrier 2 can be adjusted to be set to different predetermined angles.

In some embodiments, as shown in FIGS. 2 and 3 , the acquisition system 4 includes an acceleration sensor 401 , and the acceleration sensor 401 is attached near the installation position of the spacecraft component 6 on the back of the carrier 2 , so that the acquisition position is consistent with the spacecraft The physical parameters of the vibration response at the installation position of the component 6 are close enough, but should not interfere with each other such as collision with the spacecraft components, so as to affect the normal measurement of the acceleration sensor 401 .

The second aspect of the present invention also proposes a method for testing the pyrotechnic impact resistance of spacecraft components, through which the pyrotechnic impact environmental strength of spacecraft components can be quantitatively described, and a schematic diagram of the steps is shown in FIG. 5 . .

According to the method for testing the pyrotechnic shock resistance of a spacecraft component according to the embodiment of the second aspect of the present invention, the apparatus 1000 for testing the pyrotechnic impact resistance of a spacecraft component according to any embodiment of the first aspect of the present invention is adopted. . The method includes the following steps:

S1: Install the spacecraft component 6 to be tested at the specified installation position on the carrier 2 in a pasting manner; it is worth noting that the installation position of the spacecraft component 6 is usually located on the other side of the carrier 2 away from the laser 301 One surface is close to the incident position of the laser 301 on the carrier 2 to increase the magnitude of the laser impact, but it should not be limited to times, and the position can be adjusted according to the actual situation; at the incident position of the laser 301 on the surface of the carrier 2 An absorbing coating 202 and a transparent constraining layer 203 are added;

S2: Start the shock source 3, and perform one or more laser shocks on the carrier 2 according to actual needs;

S3: Use the acquisition system 4 to collect the physical parameters of the vibration response of a point near the installation position of the spacecraft component 6 after being impacted by the laser 301, and determine the impact intensity of the laser 301 on the spacecraft component 6 to be tested; at this time, collect The acquisition position of system 4 should be as close as possible to the installation position of spacecraft components, so that the acquisition position and the physical parameters of vibration response of the installation position of spacecraft components 6 are close enough;

S4: If the laser shock strength does not meet the actual pyrotechnic shock environment test requirements, the output energy of the shock source 3 and the installation position of the spacecraft components 6 can be adjusted, and then steps S1 to S4 are repeated to make the laser shock strength finally meet the requirements of the spacecraft components. The actual pyrotechnic shock environment test requirements of the device;

S5: Inspect the normal use performance of the spacecraft components to be tested after laser shock, so as to test the pyrotechnic impact resistance of the spacecraft components to be tested.

To sum up, steps S1 to S4 are the simulation process of the actual pyrotechnic impact environment of the spacecraft components to be tested; Impact capability test.

According to the method for testing the pyrotechnic impact resistance of spacecraft components according to the embodiment of the second aspect of the present invention, the operation method is simple and convenient. Test of pyrotechnic impact capability.

In some embodiments, the physical parameters include an acceleration time domain response signal after a point near the installation position of the spacecraft component 6 is impacted by a shock wave; after the acceleration time domain response signal is obtained, the acceleration time domain response signal will be analyzed, Draw the shock response spectrum; among them, the shock response spectrum is to apply the acceleration time domain response signal at a point near the installation position of the spacecraft component 6 to a series of single-degree-of-freedom mass-spring-damper systems with different natural frequencies, and then apply the The maximum value of the response of each single-degree-of-freedom mass-spring-damper system is plotted as the function value corresponding to different frequencies in the frequency domain. In this embodiment, the response time-domain peak value of the acceleration time-domain response signal is used as the shock-response amplitude value of the shock-response spectrum of the predetermined frequency, and then the natural frequency is valued according to a predetermined multiple, and finally the response condition of the acceleration signal is obtained. Shock Response Spectrum.

In some embodiments, the method further includes: using the same shock source 3 to carry out multiple repeated experiments on the carrier 2 to measure the vibration response of a point near the installation position of the spacecraft component 6 for exploring multiple test experiments repeatable and controllable performance.

Among them, in the above-mentioned repeated test experiment, the following steps are included: collecting the time-domain response signal of the laser 301 impact acceleration tested for multiple times; Time-domain response signal of secondary laser 301 impact acceleration; draw the single-shot laser 301 impact acceleration time-domain response signal into a shock response spectrum; convert the shock response spectrum into an image representing the shock frequency with pixel values; convert the image into a shock response spectrum ; Compare the similarities and differences of the shock response spectrum under different shock times to determine the repeatability and controllability of the test results. Among them, the shock response spectrum is when the shock of shock source 3 is applied to a series of linear, single-degree-of-freedom mass-spring-damper systems, the maximum acceleration time-domain response value of each single-degree-of-freedom system in the response motion is taken as the corresponding value of the system. The curve drawn as a function of natural frequency is shown in Figure 7. The motion model of the single-degree-of-freedom system is shown in Figure 8, and its motion differential equation is as follows:

Among them, x is the mass displacement of the single-degree-of-freedom system, and y is the base displacement.

The relative displacement z is defined as:

z=x-y (2)

Then formula (1) can be expressed as:

For the convenience of analysis, the following transformations are made:

Among them, ζ is the damping coefficient, usually obtained by the amplification factor Q, usually Q is 10, and _wn is the natural frequency.

Substituting formula (4) and formula (5) into formula (3), the following equation can be obtained:

Using the convolution integration method to solve it, the acceleration time domain response of the mass block can be obtained as follows:

Taking the time-domain peak value of the acceleration time-domain response as the amplitude of the shock response spectrum of the corresponding frequency _wn , and taking the value of the natural frequency according to the 1/12 octave frequency band, the shock response spectrum of the response can be finally obtained. In order to further evaluate the attenuation characteristics of the laser 301 shock and the influence of each parameter on the shock response, this example uses the normalized relative mean E _r and the relative maximum value M _r of the shock response spectrum, which are expressed as:

where f ₀ and f _N are the lower and upper limits of the frequency analysis range, SRS _a (f) is the analyzed impulse response spectrum, and SRS _b (f) is the reference impulse response spectrum. FIG. 9 is the data of three repetitions of the shock response spectrum measured by the device 1000 for testing the pyrotechnic shock resistance of spacecraft components according to the present invention, and the energy of the laser 301 is 1.5J at this time. It can be seen from FIG. 9 that the typical test data of the shock response spectrum obtained by the present invention has good repeatability and controllability. In addition, it can also be seen from Fig. 9 that in the frequency band of 10 ² to 10 ⁴ Hz, the amplitude of the impulse response increases with the increase of frequency, and there is almost no inflection point, which confirms that the impulse response of Laser 301 has excellent high-frequency characteristics At this time, the magnitude of the laser 301 shock response spectrum has reached the level of 200g, which has met the requirements of the mid- and far-field pyrotechnic shock test of some spacecraft components. FIG. 10 is a schematic diagram of the law of influence of the energy of the laser 301 , and it can be seen that the influence of the energy density of the laser 301 on the amplitude of the shock response spectrum is an approximate linear relationship. From Fig. 10, it can be judged that the relative coefficient of the shock response spectrum changes with the energy of the laser 301. When the single pulse energy of the laser 301 reaches several tens of J, the magnitude of the shock response generated can reach the magnitude of the near field of most pyrotechnic shock responses. Grade, that is, more than 8000g, also shows that the present invention has great development potential. In addition, it is worth noting that, considering the possible accidental damage of the spacecraft components 6 and related confidentiality requirements during the experiment, no actual spacecraft components were installed in this experiment.

In some embodiments, the following step is further included: performing a dynamic numerical simulation on the impact response of the spacecraft component 6 to be tested. Among them, the dynamic numerical simulation further includes: establishing the finite element model of the aerospace component 6 under the impact of the laser 301; analyzing the energy loss of the laser 301 under different conditions and the relationship between the impact pressure of the laser 301 and time; simulating the spacecraft components 6 Transient response under laser 301 shock. Fig. 11 and Fig. 12 are the front and side schematic diagrams of the finite element model constructed by the present invention, respectively. FIG. 13 is a partial enlarged view of the load-loading area of the bearing member 2 of FIG. 11 . In this example, the finite element model is in a free state, which is used to simulate free suspension conditions; the finite element model of the carrier 2 and the finite element model 902 of the spacecraft components adopt binding constraints, which are used to simulate the actual connection state of the spacecraft components; The pressure load applied to the loading part 903 of the finite element model of the bearing member 2 can be simplified as a triangular wave whose action time is two to three times the pulse time of the laser 301. The schematic diagram is shown in Figure 14; in Figure 14, the rising curve part and the falling curve part Corresponding to the plasma expansion process and cooling process under the action of laser 301, respectively, the peak pressure of the curve can be expressed as:

Among them, P is the peak pressure; α is the energy absorption coefficient of the carrier 2 to the laser 301; Z is the impedance coefficient between the carrier 2 and the confinement layer; I ₀ is the energy density of the incident laser 301; t _d is the pulse of the laser 301 time; t _l is the cooling end time of the plasma generated by the laser 301 .

According to the method for testing the pyrotechnic impact resistance of spacecraft components according to the embodiment of the second aspect of the present invention, the operation method is simple, the real pyrotechnic impact environment of the spacecraft components can be simulated, and the spacecraft components 6 The pyrotechnic shock resistance ability of the spacecraft can be tested; by obtaining the physical parameters of the vibration response of the spacecraft components under laser shock conditions, the pyrotechnic impact resistance of the spacecraft components 6 to be tested can be known.

In the description of this specification, reference to the terms "one embodiment," "some embodiments," "exemplary embodiment," "example," "specific example," or "some examples", etc., is meant to incorporate the embodiments A particular feature, structure, or characteristic described by an example or example is included in at least one embodiment or example of the present invention. In this specification, schematic representations of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the particular features, structures, or characteristics described may be combined in a suitable manner in any one or more embodiments or examples.

Although embodiments of the present invention have been shown and described, it will be understood by those of ordinary skill in the art that various changes, modifications, substitutions and alterations can be made in these embodiments without departing from the principles and spirit of the invention, The scope of the invention is defined by the claims and their equivalents.

Claims (14)

1. A device for testing fire impact resistance of spacecraft components is characterized by comprising:

testing the bracket;

the bearing part is freely suspended on the test support and used for bearing a spacecraft component to be tested;

the impact source is used for providing laser to generate laser impact on one surface, facing the laser, of the bearing piece, and the laser impact is transmitted to the installation position of the spacecraft component arranged on the other surface of the bearing piece through the bearing piece, so that the simulation of the real fire impact environment of the spacecraft component is realized;

and the acquisition system is connected with the bearing piece and is used for acquiring the physical parameters of vibration response of one point of the attachment of the installation position where the spacecraft component is positioned under the action of the laser shock.

2. The device for testing the fire impact resistance of a spacecraft component as recited in claim 1, further comprising a light path control system, said light path control system being disposed between said impact source and said carrier for controlling a laser conduction light path.

3. The device for testing the fire impact resistance of the spacecraft component as recited in claim 2, wherein the optical path control system comprises a reflector and a focusing lens, and the laser emitted by the impact source is reflected to the focusing lens through the reflector and then focused by the focusing lens to irradiate on the bearing member.

4. The apparatus for testing the fire impact resistance of a spacecraft component as recited in claim 3, wherein a focal length of the focusing lens is set to be adjustable.

5. The device for testing the fire impact resistance of the spacecraft component according to any one of claims 1 to 4, wherein the bearing piece comprises a base layer, an absorption coating layer arranged on the side of the base layer facing the laser, and a transparent constraint layer arranged on the absorption coating layer;

when the laser penetrates through the transparent constraint layer and reaches the absorption coating, the absorption coating absorbs energy of the laser to generate plasma, the plasma absorbs residual energy of the laser under the constraint action of the transparent constraint layer and expands rapidly, shock waves are generated and transmitted to the installation position of the spacecraft component fixed on the other surface of the bearing component through the bearing component, and vibration impact is caused on the spacecraft component.

6. The device for testing the fire impact resistance of the spacecraft component as recited in any one of claims 1 to 4, further comprising a flexible member, wherein one end of the flexible member is fixed on the test support, and the other end of the flexible member is fixed on the bearing member, so that the bearing member is freely suspended on the test support.

7. An apparatus for testing fire impact resistance of a spacecraft component as claimed in any one of claims 1 to 4, wherein the incidence angle of the laser light relative to the carrier is set at an adjustable predetermined angle.

8. The device for testing the fire impact resistance of the spacecraft component as recited in any one of claims 1 to 4, wherein the collection system comprises an acceleration sensor, and the acceleration sensor is attached near the installation position of the spacecraft component of the bearing component, but should not collide with the spacecraft component and interfere with the spacecraft component.

9. A method for testing the fire impact resistance of a spacecraft component, which is characterized in that the device for testing the fire impact resistance of the spacecraft component, which is disclosed by any one of claims 1 to 8, is adopted, and comprises the following steps:

s1: mounting a spacecraft component to be tested at a specified mounting position on the bearing piece in a pasting manner;

s2: starting the impact source, and performing one or more times of laser impact on the bearing piece according to actual requirements;

s3: acquiring physical parameters of vibration response of a point near the installation position of the spacecraft component under the action of the laser shock by using the acquisition system, and determining the laser shock strength borne by the spacecraft component to be tested;

s4: if the laser impact strength does not meet the test requirement of the actual fire impact environment, adjusting the output energy of the impact source and the installation position of the spacecraft component; then, the steps S1 to S4 are repeatedly executed, so that the laser impact strength finally meets the requirement of testing the actual fire impact environment of the spacecraft component;

and S5, checking the normal use performance of the spacecraft component after the laser impact, thereby realizing the test of the fire impact resistance of the spacecraft component to be tested.

10. The method for testing the fire impact resistance of an aerospace device according to claim 9,

the physical parameters comprise acceleration time domain response signals of a point near the installation position of the spacecraft component after the point is impacted by the laser.

11. The method for testing the fire impact resistance of a spacecraft component as recited in claim 10, further comprising the steps of:

analyzing the strength of the acceleration time domain response signal of a point near the installation position of the spacecraft component after the acceleration time domain response signal is subjected to the laser shock action, and drawing a shock response spectrum;

in the shock response spectrum, the response time domain peak value of the acceleration time domain response signal is used as the shock response spectrum amplitude value of the shock response spectrum with the preset frequency.

12. The method for testing the fire impact resistance of a spacecraft component as recited in claim 11, further comprising the steps of: and the repeatability and controllability of the test result are checked by adopting the same impact source.

13. The method for testing the fire impact resistance of a spacecraft component as recited in claim 12,

collecting laser shock acceleration time domain response signals tested for multiple times;

decomposing the laser shock acceleration time domain response signals tested for multiple times into single laser shock acceleration time domain response signals according to a time sequence;

drawing the single laser shock acceleration time domain response signal into the shock response spectrum;

the impulse response spectrum is converted into an image representing the impulse frequency in pixel values.

14. The method for testing the fire impact resistance of the spacecraft component as recited in claim 9, further comprising the steps of:

1) establishing a laser shock finite element model of the spacecraft component;

2) and carrying out dynamic numerical simulation on the laser shock response of the spacecraft component.

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