CN114492142B - Device and method for testing fire impact resistance of spacecraft component - Google Patents
Device and method for testing fire impact resistance of spacecraft component Download PDFInfo
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- CN114492142B CN114492142B CN202210179197.4A CN202210179197A CN114492142B CN 114492142 B CN114492142 B CN 114492142B CN 202210179197 A CN202210179197 A CN 202210179197A CN 114492142 B CN114492142 B CN 114492142B
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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 support and 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 physical parameters of vibration response of a point near the installation position of the spacecraft component under the impact of 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
Technical Field
The invention relates to the technical field of simulation of an explosive impact environment of spacecraft components and testing of an explosive impact resistance, and mainly relates to a device and a method for testing the explosive impact resistance of the spacecraft components.
Background
The impact environment of the aerospace initiating explosive device is structural impact response caused by explosive separation of initiating explosive devices in working processes of satellite-rocket separation, fairing separation, interstage separation, component unfolding and the like. The impact environment of the aerospace firer has the characteristics of high frequency, transient state and high magnitude, and is very easy to damage precise electronic components and other instruments and equipment on the spacecraft. The flame impact environment of the spaceflight is one of the most harsh mechanical environments experienced by the spaceflight in the whole life cycle, and in the development process of the spaceflight, a flame impact simulation test needs to be carried out on the ground in advance, namely, the flame impact resistance of a spacecraft element is examined and checked. Up to now, the simulation experiment method for anti-firer impact capability is mainly divided into two types of firer explosion type and non-firer explosion type.
The former directly adopts gunpowder or a detonator, and is loaded on different types of resonance devices in a certain mode, and the initiating explosive impact response is obtained by igniting the gunpowder or the detonator. The simulation mode belongs to destructive excitation tests, has poor operability and result repeatability, can obtain an impact response test result with high relative precision by performing multiple trial and error before formal tests, has high test cost and poor safety, and is easy to cause great damage to testers and test devices. The latter mainly comprises two types, namely a mechanical vibration type and a vibration table type, wherein a pendulum bob, an air gun or a vibration table is respectively used as an impact generator, and the vibration response of the vibration generator is obtained through a resonance device acted by structures such as a resonance plate, a resonance rod and the like so as to simulate the fire impact response of a spacecraft element under a real condition.
The non-explosive test method has good operability and repeatability and low test cost, but has a narrow frequency spectrum range (the frequency spectrum range of a vibration table type can only reach about 3000 Hz), and small high-frequency response amplitude, and cannot simulate the high-frequency response characteristics of explosive impact (the main frequency of the high-frequency response amplitude is between 100Hz and 100 Khz).
Disclosure of Invention
The present invention is directed to solving at least one of the problems of the prior art. Therefore, the invention aims to provide a device for testing the fire impact resistance of a spacecraft component, which can simulate the fire impact environment of the spacecraft and can test the fire impact resistance of the spacecraft component.
The device for testing the fire impact resistance of the spacecraft component according to the embodiment of the first aspect of the invention comprises:
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 physical parameters of vibration response of one point of an attachment at the installation position where the spacecraft component is positioned under the action of the laser shock.
The device for testing the fire impact resistance of the spacecraft component in the embodiment of the first aspect of the invention has the advantages of simple structure and simple and convenient operation method, can well represent the high-frequency fire impact characteristic, and realizes the simulation of the real fire impact environment of the spacecraft component so as to test the fire impact resistance of the spacecraft component.
In some embodiments, the laser beam path control system is disposed between the impact source and the bearing member for controlling the laser beam path.
In some embodiments, the optical path control system includes a reflector and a focusing lens, and the laser emitted by the impact source is reflected by the reflector to the focusing lens, and then is focused by the focusing lens to irradiate on the bearing member.
In some embodiments, the focal length of the focusing lens is set to be adjustable.
In some embodiments, the carrier comprises a base layer, an absorptive coating disposed on a side of the base layer facing the laser, and a transparent constraining layer disposed on the absorptive coating;
when the laser penetrates through the transparent constraint layer and then reaches the absorption coating, the absorption coating absorbs the energy of the laser and generates plasma, the plasma absorbs the residual energy of the laser to rapidly expand (namely, the plasma explosion phenomenon) under the constraint action of the transparent constraint layer, shock waves are generated and transmitted to the installation position of the spacecraft component to be tested, which is fixed on the other surface of the bearing component, through the bearing component, and therefore the simulation of the initiating explosive impact environment of the spacecraft component is realized.
In some embodiments, the test device further comprises a flexible member, one end of the flexible member is fixed to the test rack, 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 rack.
In some embodiments, the angle of incidence of the laser light relative to the carrier is set at an adjustable predetermined angle.
In some embodiments, the acquisition system includes an acceleration sensor attached near the location where the carrier spacecraft components are mounted, but not interfering with the spacecraft components.
The invention also provides a method for testing the fire work impact resistance of the spacecraft component.
According to the method for testing the fire impact resistance of the spacecraft component in the embodiment of the second aspect of the invention, the device for testing the fire impact resistance of the spacecraft component in the embodiment of the first aspect of the invention comprises the following steps:
s1: mounting a spacecraft component to be tested at a specified mounting position on the bearing piece in a sticking manner;
s2, starting the impact source, and carrying out 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 test requirement of 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.
In conclusion, the steps S1 to S4 are the simulation process of the real fire impact environment of the spacecraft component to be tested; and S5, testing the fire impact resistance of the spacecraft component in the fire impact environment determined in the steps S1 to S4.
According to the method for testing the fire impact resistance of the spacecraft component, the operation method is simple and convenient, and the test of the fire impact resistance of the spacecraft component is realized by simulating the real fire impact environment of the spacecraft component.
In some embodiments, the physical parameter comprises an acceleration time domain response signal of a point near the installation position of the spacecraft component after being impacted by the laser.
In some embodiments, the method further comprises the steps of:
analyzing the acceleration time domain response signal and drawing an impact response spectrum;
and 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 with the preset frequency.
In some embodiments, the method further comprises the steps of: and the repeatability and controllability of the test result are checked by adopting the same impact source.
In some embodiments, laser shock acceleration time domain response signals of multiple tests are collected;
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;
converting the impulse response spectrum into an image representing the impulse frequency in pixel values;
and converting the image into an impact response spectrogram.
In some embodiments, the method further comprises the steps of: establishing a laser shock finite element model of the spacecraft component; and carrying out dynamic numerical simulation on the impact response of the spacecraft component.
Additional aspects and advantages of the invention will be set forth in part in the description which follows and, in part, will be obvious from the description, or may be learned by practice of the invention.
Drawings
The above and/or additional aspects and advantages of the present invention will become apparent and readily appreciated from the following description of the embodiments, taken in conjunction with the accompanying drawings of which:
fig. 1 is a schematic diagram of an apparatus for testing fire impact resistance of a spacecraft component according to an embodiment of the first aspect of the invention.
Fig. 2 is an enlarged view of a portion of the carrier region of fig. 1.
Fig. 3 is a schematic diagram of another apparatus for testing the fire impact resistance of a spacecraft component according to an embodiment of the first aspect of the invention.
Fig. 4 is an enlarged view of a portion of the carrier region of fig. 3.
Fig. 5 is a schematic diagram of a method for testing the fire impact resistance of a spacecraft component according to a second aspect of the embodiment of the invention.
Fig. 6a to 6c are schematic diagrams illustrating different included angles of rotation of a bearing component in a method for testing the fire impact resistance of a spacecraft component according to a second aspect of the present invention.
Fig. 7 is a schematic diagram of an impulse response spectrum of an impulse environment of a fire power impact obtained by the apparatus for testing the impulse resistance of a spacecraft component under laser impact according to the first aspect of the present invention in the method for testing the impulse resistance of a spacecraft component according to the second aspect of the present invention.
FIG. 8 is a schematic diagram of a single degree of freedom mass-spring-damping system dynamics model required to synthesize the impact response spectrum shown in FIG. 6.
FIG. 9 is an impulse response spectrum measured by the apparatus of FIG. 1.
FIG. 10 is a graph showing the effect of laser fluence on the amplitude versus coefficient of the impulse response spectrum measured by the apparatus of FIG. 1.
Fig. 11 is a schematic front view of a finite element model in a method for testing the fire impact resistance of a spacecraft component according to a second aspect of the present invention.
Fig. 12 is a schematic side view of a finite element model in a method for testing the fire impact resistance of a spacecraft component according to a second aspect of the present invention.
Fig. 13 is an enlarged view of a portion of the load carrying region of the load carrying member of fig. 11.
FIG. 14 is a pressure load diagram of a finite element model load loading location of a load bearing member.
Reference numerals:
Transparent constraining layer 203 of base layer 201 absorptive coating 202 of carrier 2
Impact source 3 laser 301
Optical path control system 5 mirror 501 focusing lens 502
Carrier finite element model 901
Spacecraft component finite element model 902
Load bearing finite element model load loading part 903
Detailed Description
Reference will now be made in detail to embodiments of the present invention, examples of which are illustrated in the accompanying drawings, wherein like or similar reference numerals refer to the same or similar elements or elements having the same or similar function throughout. The embodiments described below with reference to the accompanying drawings are illustrative only for the purpose of explaining the present invention, and are not to be construed as limiting the present invention.
The device 1000 and the method for testing the fire impact resistance of the spacecraft component according to the embodiment of the invention are described below with reference to fig. 1 to 14.
As shown in fig. 1 to 4, the apparatus 1000 for testing the fire impact resistance of a spacecraft component according to the embodiment of the first aspect of the present invention can simulate the environment of the fire impact of the spacecraft and test the fire impact resistance of the spacecraft component 6.
The device 1000 for testing the fire impact resistance of the spacecraft component according to the embodiment of the first aspect of the invention comprises a test support 1, a bearing part 2, an impact source 3 and an acquisition system 4. The test device comprises a test support 1, a bearing piece 2, a test device component 6 and a test device, wherein the bearing piece 2 is freely suspended on the test support 1, and the bearing piece 2 is used for bearing the spacecraft component 6 to be tested; the impact source 3 is used for providing laser 301 to generate laser impact on the surface, facing the laser, of the bearing piece 2, and the laser impact is transmitted to the installation position of the spacecraft component 6 set on the other surface of the bearing piece 2 through the bearing piece 2, so that simulation of a real firer impact environment of the spacecraft component 6 is realized; the acquisition system 4 is connected with the carrier 2 and used for acquiring the physical parameters of the vibration response of a point near the installation position of the spacecraft component 6 after the impact action of the laser 301.
Specifically, the test rack 1 is a rack structure to hang the fixed carrier 2 freely, so that the carrier 2 is suspended (for example, as shown in fig. 1 and fig. 2) to facilitate the experiment.
The bearing part 2 is freely suspended on the test support 1, and the bearing part 2 is used for bearing a spacecraft component 6 to be tested; that is to say, on the one hand, the bearing part 2 is fixed on the test support 1 in a suspended state in a free suspension mode, and on the other hand, the bearing part 2 is used for being fixed with the spacecraft component 6 to be tested, so that the simulation of the free boundary of the initiating explosive impact of the spacecraft component is realized.
The impact source 3 is used for providing laser 301 to generate laser impact on the surface of the bearing piece 2 facing the laser, and the laser impact is transmitted to the installation position of the spacecraft component 6 fixed on the other surface of the bearing piece 2 through the bearing piece 2, so that the simulation of the real fire impact environment of the spacecraft component 6 is realized. Specifically, the impact source 3 may be a laser generator, and the structure of the laser generator includes a power supply system, a laser generator, a cooling system, and the like, and preferably, the laser generator is a solid laser generator capable of generating laser 301 with high energy density and narrow pulse width; when the laser 301 provided by the impact source 3 acts on one surface of the bearing piece 2 facing the laser, a plasma explosion phenomenon occurs on the surface of the bearing piece 2 (i.e., the surface facing the laser), and shock waves are generated and transmitted to the installation position of the spacecraft component 6 fixed on the other surface of the bearing piece 2 (the surface facing away from the laser) through the bearing piece 2, so that the spacecraft component 6 generates a high-frequency vibration response very close to that under a real firer impact condition, and the simulation of a real firer impact environment where the spacecraft component 6 is located is realized.
The acquisition system 4 is connected with the bearing part 2 and is used for acquiring the physical parameters of the vibration response of a point near the installation position of the spacecraft component 6 under the impact action of the laser 301. Since the installation position of the spacecraft component 6 is provided with equipment, the physical parameters cannot be directly measured, and in the actual measurement process, a point near the installation position of the spacecraft component 6 is generally used for approximate replacement. By analyzing and processing the physical parameters, the intensity of the fire impact environment where the spacecraft component 6 is located can be obtained, and the testing of the fire impact resistance of the spacecraft component 6 is realized by testing the normal use performance of the spacecraft component 6 after laser impact.
When the device 1000 for testing the fire impact resistance of the spacecraft component is used for testing the spacecraft component 6, the impact source 3 can be placed on the workbench 8, and the test support 1 is arranged on one side of the impact source 3; fixing a spacecraft component 6 to be tested at a specified mounting position on the bearing member 2; starting the impact source 3 to perform one or more times of laser impact on the bearing piece 2; acquiring physical parameters of vibration response of a point near the installation position of the spacecraft component 6 under the action of laser impact by using an acquisition system 4; and (4) inspecting the normal use performance of the spacecraft component 6 to be tested after laser impact.
The device 1000 for testing the fire impact resistance of the spacecraft component in the embodiment of the first aspect of the invention has the advantages of simple structure and simple and convenient operation method, can well represent the high-frequency fire impact characteristic, realizes the simulation of the real fire impact environment of the spacecraft component, and is convenient for testing the fire impact resistance of the spacecraft component.
In some embodiments, as shown in fig. 1 and 3, a light path control system 5 is further included, the light path control system 5 being disposed between the impact source 3 and the carrier 2 for control of the laser conduction light path. That is to say, by arranging the optical path control system 5, after the laser 301 emitted by the impact source 3 is focused, the energy of the laser 301 is more concentrated, and it can be effectively ensured that the focused laser 301 can generate impact waves of the magnitude required by the experiment when acting on the bearing member 2. It should be noted that the optical path control system 5 is generally placed on the table 8, but is not limited thereto and may be adjusted according to actual conditions.
In some embodiments, as shown in fig. 3, the optical path control system 5 includes a mirror 501 and a focusing lens 502, and the laser light 301 emitted by the impact source 3 is reflected by the mirror 501 to the focusing lens 502, and then focused by the focusing lens 502 to irradiate on the carrier 2. The mirror 501 is used for changing the propagation path of the laser 301 emitted by the impact source 3, and the focusing lens 502 is used for adjusting the spot diameter of the laser 301 emitted by the impact source 3.
In some embodiments, the focal length of the focusing lens 502 is set to be adjustable. In this way, the focal length of the focusing lens 502 can be set according to experimental requirements to adjust the size of the light spot.
In some embodiments, as shown in fig. 2 and 4, the carrier 2 comprises a substrate layer 201, an absorptive coating 202 disposed on a side of the substrate layer 201 facing the laser 301, and a transparent constraining layer 203 disposed on the absorptive coating 202; when the laser light 301 reaches the absorption coating 202 after passing through the transparent constraining layer 203, the absorption coating 202 absorbs the energy of the laser light 301 and generates plasma, and the plasma rapidly expands (i.e., plasma explosion phenomenon) after absorbing the remaining energy of the laser light 301 under the constraint action of the transparent constraining layer 203, so as to generate shock waves and transmit the shock waves to the spacecraft component 6 fixed on the other surface of the carrier 2 through the carrier 2.
Specifically, an absorption coating 202 and a transparent constraint layer 203 are arranged on one side of the substrate layer 201 facing the laser 301, wherein the absorption coating 202 is located between the substrate layer 201 and the transparent constraint layer 203, and the other side of the substrate layer 201 can fix the spacecraft component 6 to be tested. The provision of transparent confinement layer 203 and absorptive coating 202 can effectively enhance the laser 301 impact pressure and extend the duration of the laser 301 impact. After being focused by the optical path control system 5, the laser 301 emitted by the impact source 3 first passes through the transparent constraint layer 203, and since the transparent constraint layer 203 is transparent to the laser 301, the laser 301 can reach the absorption coating 202 after penetrating through the transparent constraint layer 203. Under the ionization action of the high-energy-density laser 301, the absorption coating 202 absorbs the energy of the laser 301 and generates plasma, and the plasma absorbs the residual energy of the laser 301 and rapidly expands (i.e., plasma explosion phenomenon) under the constraint action of the transparent constraint layer 203, so as to generate shock waves and transmit the shock waves to the spacecraft component 6 fixed on the other surface of the carrier 2 through the carrier 2. The installation position of the spacecraft component 6 is usually located on the surface of the carrier 2 opposite to the laser 301 and close to the incident position of the laser 301, so as to improve the laser shock magnitude, but not limited to this, and can be adjusted according to actual requirements.
Optionally, the substrate layer 201 may be made of a material such as an aluminum alloy, a titanium alloy, or stainless steel, but is not limited thereto, and may be adjusted according to actual requirements.
The absorptive coating 202 may be aluminum foil or black tape positioned on the side of the substrate 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 0.1mm, but not limited to this, and may be adjusted according to actual requirements.
Transparent constraining layer 203 may be distilled water or transparent glass. The impact effect of the distilled water enhanced laser 301 is poorer than that of transparent glass, but the distilled water used as the transparent constraint layer 203 has the advantages of convenience in use, good experimental result repeatability and the like. Therefore, the material of the transparent constraining layer 203 is preferably distilled water, and the structural schematic diagram thereof is shown in fig. 1. In addition, when the transparent constraining layer 203 is made of distilled water, the testing apparatus further includes a water spraying pipe externally connected to the water source, and the distilled water is sprayed from the water spraying pipe, so as to form the transparent constraining layer 203 with a certain thickness on the supporting member 2. In addition, it should be noted that when the transparent constraining layer 203 is distilled water, a portion of the distilled water near the absorbing coating 202 is ionized by the laser 301.
In some embodiments, as shown in fig. 1, the testing device further comprises a flexible member 7, one end of the flexible member 7 is fixed on the testing support 1, and the other end is fixed on the carrying member 2, so that the carrying member 2 is freely suspended from the testing support 1. Therefore, the carrier 2 is fixed on the test support 1 by the flexible piece 7, the assembly operation is very simple, and the position of the carrier can be adjusted. For example, as shown in fig. 1, the carrier 2 can be freely suspended from the test rack 1 by four flexible members 7, and the height of the carrier 2 can be adjusted by changing the lengths of the flexible members 7. In addition, it should be noted that the number of the flexible members 7 is not limited to four, and may be selected according to actual requirements.
In some embodiments, the angle of incidence of the laser light 301 with respect to the carrier 2 is set at an adjustable predetermined angle. The angle of incidence may be understood as the angle between the laser light 301 and the carrier 2. Changing the included angle between the laser 301 and the carrier 2 affects the spot shape of the laser 301 and the energy distribution of the laser 301, and thus affects the shock wave impact pressure of the laser 301. Therefore, the included angle between the laser 301 and the bearing member 2 is set to be different preset angles, physical parameters of vibration response after the mounting position of the spacecraft component to be tested is subjected to the impact action of the laser 301 under different preset angles (as shown in fig. 6a to 6 c) are collected, the influence rule of the included angle between the laser 301 and the bearing member 2 on the fire impact environment of the mounting position of the spacecraft component 6 is quantitatively determined, and the fire impact environment of the spacecraft component 6 to be tested is accurately adjusted. Wherein, by rotating the angle of the carrier 2 (as shown in fig. 6a 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 fig. 2 and 3, the collection system 4 includes an acceleration sensor 401, and the acceleration sensor 401 is attached near the mounting position of the spacecraft component 6 on the back of the carrier 2, so that the vibration response physical parameters of the collection position and the mounting position of the spacecraft component 6 are close enough, but should not interfere with each other such as collision with the spacecraft component, so as to affect the normal measurement of the acceleration sensor 401.
The invention also provides a method for testing the fire impact resistance of the spacecraft component, the intensity of the fire impact environment of the spacecraft component can be quantitatively described through the method, and the step schematic diagram is shown in fig. 5.
According to the method for testing the fire impact resistance of the spacecraft component in the embodiment of the second aspect of the invention, the device 1000 for testing the fire impact resistance of the spacecraft component in any one of the embodiments of the first aspect of the invention is adopted. The method comprises the following steps:
s1, mounting a spacecraft component 6 to be tested at a specified mounting position on a bearing part 2 in a pasting manner; it should be noted that the mounting position of the spacecraft component 6 is usually located on the other surface of the carrier 2 opposite to the laser 301 and close to the incident position of the laser 301 on the carrier 2 to increase the laser impact magnitude, but should not be limited to the second time, and the position can be adjusted according to actual situations; an absorption coating 202 and a transparent constraint layer 203 are added at the incident position of the laser 301 on the surface of the carrier 2;
s2: starting the impact source 3, and performing one or more times of laser impact on the bearing piece 2 according to actual requirements;
s3: acquiring physical parameters of vibration response of a point near the installation position of the spacecraft component 6 under the impact action of the laser 301 by using the acquisition system 4, and determining the impact strength of the laser 301 borne by the spacecraft component 6 to be tested; at the moment, the distance between the acquisition position of the acquisition system 4 and the installation position of the spacecraft component is close to the maximum extent, so that the vibration response physical parameters of the acquisition position and the installation position of the spacecraft component 6 are close enough;
s4: if the laser impact strength does not meet the actual testing requirement of the fire impact environment, the output energy of the impact source 3 and the installation position of the spacecraft component 6 can be adjusted, and then the steps S1 to S4 are repeatedly executed, so that the laser impact strength finally meets the actual testing requirement of the fire impact environment of the spacecraft component;
and S5, testing the normal use performance of the spacecraft component to be tested after laser shock, thereby realizing the test of the fire shock resistance of the spacecraft component to be tested.
In conclusion, the steps S1 to S4 are the simulation process of the real fire impact environment of the spacecraft component to be tested; and step S5, testing the fire impact resistance of the spacecraft component 6 under the fire impact environment determined in the steps S1 to S4.
According to the method for testing the fire impact resistance of the spacecraft component in the embodiment of the second aspect of the invention, the operation method is simple and convenient, and the test of the fire impact resistance of the spacecraft component 6 is realized by simulating the real fire impact environment of the spacecraft component.
In some embodiments, the physical parameters include an acceleration time domain response signal after a shock wave is applied to a point near the installation position of the spacecraft component 6; after the acceleration time domain response signal is obtained, analyzing the acceleration time domain response signal, and drawing an impact response spectrum; the impact response spectrum is a frequency domain curve obtained by applying an acceleration time domain response signal of a point near the installation position of the spacecraft component 6 to a series of single-degree-of-freedom mass-spring-damping systems with different natural frequencies and then drawing the maximum value of the response of each single-degree-of-freedom mass-spring-damping system as function values corresponding to different frequencies. In the embodiment, the response time domain peak value of the acceleration time domain response signal is used as the impact response amplitude of the impact response spectrum with the preset frequency, and then the natural frequency is subjected to value taking according to the preset multiple, so that the impact response spectrum under the acceleration signal response condition is finally obtained.
In some embodiments, further comprising: the same impact source 3 is adopted to carry out repeated experiments on the bearing part 2, and the vibration response of a point near the installation position of the spacecraft component 6 is measured, so that the repeatable performance and the controllability of the repeated testing experiments are researched.
The method comprises the following steps of: collecting laser 301 impact acceleration time domain response signals tested for multiple times; decomposing the laser 301 impact acceleration time domain response signals tested for multiple times into single laser 301 impact acceleration time domain response signals according to a time sequence; drawing a single laser 301 shock acceleration time domain response signal into a shock response spectrum; converting the impulse response spectrum into an image representing the impulse frequency in pixel values; converting the image into an impact response spectrogram; and comparing the differences of the impact response spectrums under different impact times, and determining the repeatability and controllability of the test result. The impulse response spectrum is a curve that plots the maximum acceleration time-domain response value in the response motion of each single-degree-of-freedom system as a function corresponding to the natural frequency of the system when the impulse of the impulse source 3 is applied to a series of linear, single-degree-of-freedom mass-spring-damper systems, and a schematic diagram thereof is shown in fig. 7. The motion model of the single-degree-of-freedom system is shown in fig. 8, and the motion differential equation is as follows:
wherein, x is the displacement of the mass block of the single-degree-of-freedom system, and y is the basic displacement.
Defining the relative displacement z as:
z＝x-y (2)
equation (1) can be expressed as:
for convenient analysis, the following transformations were made:
where ζ is the damping coefficient, and is usually determined by an amplification factor Q, usually Q is 10 n Is the natural frequency.
Substituting equation (4) and equation (5) into equation (3) can obtain the following equations:
solving the mass block acceleration time domain response by adopting a convolution integral method, wherein the mass block acceleration time domain response can be obtained as follows:
taking the acceleration time domain response time domain peak value as the corresponding frequency w n The amplitude of the impulse response spectrum is subjected to value taking on the natural frequency according to 1/12 octave, and finally the impulse response spectrum of the response is obtained. In order to further evaluate the attenuation characteristics of the laser 301 impact and the influence rule of each parameter on the impact response, the normalized relative mean value E of the impact response spectrum is adopted in the embodiment r And the relative maximum value M r Respectively expressed as:
wherein f is 0 And f N Is the lower and upper limit of the frequency analysis range, SRS a (f) Is an analyzed shock response spectrum, SRS b (f) Is a reference impulse response spectrum. Fig. 9 is three times of repeated test data of an impact response spectrum measured by the device 1000 for testing the fire impact resistance of a spacecraft component under the same working condition, wherein the energy of laser 301 is 1.5J. As can be seen from FIG. 9, the typical test data of the impact response spectrum obtained by the present invention has good repeatability and controllability. In addition, as can also be seen in FIG. 9, at 10 2 ～10 4 Within the Hz frequency range, the impact response amplitude is increased along with the increase of the frequency, and an inflection point hardly appears, thereby proving thatThe laser 301 impact response has excellent high-frequency characteristics, and the laser 301 impact response spectrum magnitude reaches 200g level, so that the requirements of middle and far field fire impact tests of parts of spacecraft components are met. Fig. 10 is a schematic diagram illustrating the influence rule of laser 301 energy, and it can be seen that the influence of laser 301 energy density on the amplitude of the impulse response spectrum is approximately linear. It can be judged from the trend of the relative coefficient of the impulse response spectrum of fig. 10 along with the change of the energy of the laser 301, when the single pulse energy of the laser 301 reaches dozens of J, the magnitude of the impulse response generated by the laser 301 can reach the magnitude of most of the fire impulse response near fields, namely more than 8000g, which also shows that the invention has great development potential. In addition, it is noted that, in consideration of the possible accidental damage of the spacecraft components 6 and the related privacy requirements during the experiment, no actual spacecraft components are installed in the experiment.
In some embodiments, the method further comprises the steps of: and carrying out dynamic numerical simulation on the impact response of the spacecraft component 6 to be tested. Wherein the dynamic numerical simulation further comprises: establishing a finite element model of the aerospace component 6 under the impact action of the laser 301; analyzing the energy loss of the laser 301 and the change relation of the impact pressure of the laser 301 along with time under different conditions; the transient response of the spacecraft components 6 under the impact of the laser 301 is simulated. FIGS. 11 and 12 are schematic diagrams of a front and side view of a finite element model constructed according to the present invention. Fig. 13 is an enlarged view of a portion of the load-bearing region of the carrier 2 of fig. 11. In this example, the finite element model is in a free state for simulating a free suspension condition; binding constraint is adopted between the finite element model of the bearing part 2 and the finite element model 902 of the spacecraft component, and the binding constraint is used for simulating the actual connection state of the spacecraft component; the loading pressure load applied to the load loading part 903 of the finite element model of the bearing part 2 can be simplified into a triangular wave with the action time being two to three times of the pulse time of the laser 301, and a schematic diagram of the triangular wave is shown in fig. 14; in fig. 14, the rising curve portion and the falling curve portion correspond to the plasma expansion process and the cooling process under the action of the laser 301, respectively, and the peak pressure of the curve can be expressed as:
wherein P is the peak pressure; α is an energy absorption coefficient of the carrier 2 to the laser 301; z is the impedance coefficient between the carrier 2 and the constraining layer; i is 0 The incident laser 301 energy density; t is t d Is the laser 301 pulse time; t is t l Is the end time of cooling of the plasma generated by laser 301.
According to the method for testing the fire impact resistance of the spacecraft component in the embodiment of the second aspect of the invention, the operation method is simple and convenient, the real fire impact environment of the spacecraft component can be simulated, and the fire impact resistance of the spacecraft component 6 can be tested; the physical parameters of the vibration response of the spacecraft component under the laser shock condition are obtained, so that the fire shock resistance of the spacecraft component 6 to be tested can be obtained.
In the description herein, references to the description of the term "one embodiment," "some embodiments," "an illustrative embodiment," "an example," "a specific example," or "some examples" or the like are intended to mean that a particular feature, structure, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the terms used above do not necessarily refer to the same embodiment or example. Furthermore, the particular features, structures, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples.
While embodiments of the 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 to the embodiments without departing from the principles and spirit of the invention, the scope of which is defined by the claims and their equivalents.
Claims (5)
1. A method for testing the fire impact resistance of a spacecraft component is characterized in that a device for testing the fire impact resistance of the spacecraft component is adopted, and the device for testing the fire impact resistance of the spacecraft component comprises the following steps:
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 spacecraft component generates high-frequency vibration response which is very similar to that under the real firer impact condition, and the simulation of the real firer impact environment of the spacecraft component is realized;
the acquisition system is connected with the bearing piece and is used for acquiring physical parameters of vibration response of a point near the installation position of the spacecraft component under the action of the laser impact;
the carrier comprises a base layer, an absorptive coating disposed on a side of the base layer facing the laser, and a transparent constraining layer disposed on the absorptive coating;
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 rapidly expands to generate shock waves, and the shock waves are transmitted to the installation position of the spacecraft component fixed on the other surface of the bearing piece through the bearing piece to cause vibration impact on the spacecraft component;
the method for testing the fire impact resistance of the spacecraft component 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 testing 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 test requirement of the actual fire impact environment of the spacecraft component;
s5, checking the normal use performance of the spacecraft component after the laser shock, thereby realizing the test of the fire shock resistance of the spacecraft component to be tested;
the method for testing the fire impact resistance of the spacecraft further comprises the following steps:
1) Establishing a laser impact finite element model of a spacecraft component;
2) Carrying out dynamic numerical simulation on the laser shock response of the spacecraft component;
wherein, the dynamic numerical simulation of the laser shock response of the spacecraft component comprises the following steps:
analyzing the change relation of the laser impact strength with time under different conditions, comprising the following steps:
wherein, P is the peak value of the impact strength, alpha is the energy absorption coefficient of the bearing element to the laser, and Z is the impedance coefficient between the bearing element and the constraint layer; I.C. A 0 Is the incident laser energy density.
2. The method for testing the fire impact resistance of the spacecraft component according to claim 1, wherein the physical parameters comprise an acceleration time domain response signal of a point near the installation position of the spacecraft component after the laser impact action.
3. The method for testing the fire impact resistance of the spacecraft component as recited in claim 2, further comprising the steps of:
analyzing the strength of an acceleration time domain response signal of a point near the installation position of the spacecraft component after the point is subjected to the laser shock effect, 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.
4. The method for testing the fire impact resistance of the spacecraft component as recited in claim 3, further comprising the steps of: and the repeatability and controllability of the test result are checked by adopting the same impact source.
5. The method for testing the fire impact resistance of a spacecraft component as recited in claim 4,
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.
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| CN111666700A (en) * | 2020-05-12 | 2020-09-15 | 清华大学 | Laser excitation and fire impact equivalent method |
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| DE102017219019A1 (en) * | 2017-10-25 | 2019-04-25 | Airbus Operations Gmbh | Method, use, system and mobile laser shock treatment station for laser shock blasting of a component |
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| CN109443682B (en) * | 2018-11-29 | 2024-02-09 | 清华大学 | Laser excitation type initiating explosive device impact simulation device |
| CN109387347B (en) * | 2018-11-29 | 2024-07-30 | 清华大学 | Laser excitation type initiating explosive device impact ground simulation test device |
| CN109506864A (en) * | 2018-12-12 | 2019-03-22 | 上海卫星装备研究所 | Spacecraft shock environment analogue means |
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| CN111666700A (en) * | 2020-05-12 | 2020-09-15 | 清华大学 | Laser excitation and fire impact equivalent method |
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