Explosive impact isolation method using cellular porous materials

Abstract

The present application relates to the technical field of pyrotechnic impact equipment, and more particularly to a pyrotechnic impact isolation method using honeycomb porous materials. The present application uses the complexity of the transmission path of honeycomb porous materials to construct the satellite-rocket interface, increase the stress wave impedance of the satellite-rocket interface, reduce the stress wave transmissivity, and achieve the isolation of the pyrotechnic impact response. The present application uses the explicit finite element method to model and analyze the transient dynamics of the relevant structure, and studies the pyrotechnic impact isolation characteristics of honeycomb porous materials. The present application can effectively reflect the stress wave, dissipate the impact energy, achieve the effective isolation and suppression of the pyrotechnic impact, and improve the method of traditional satellite-rocket connection.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of pyrotechnic impact device, more particularly, to a pyrotechnic impact isolation method using honeycomb porous material. BACKGROUND

[0002] In the field of aerospace engineering, explosive devices such as separation nuts and explosive bolts are widely used in the connection and separation devices of satellite launch systems. The high-frequency and transient load caused by the initiation of the pyrotechnic separation device drives the movement of the mechanism, and the impact response of the structure under the action of the load is one of the most severe mechanical environments experienced by the spacecraft. It is easy to cause damage to nearby microstructures and hardware, and even cause the failure of the space launch mission. In the traditional method, the main method to reduce the impact response of pyrotechnics is the flexible energy absorption method and the rigid isolation method. In aerospace engineering, the rigid isolation method is more widely used than the flexible energy absorption method. This method changes the wave impedance of the connection interface by changing the structure, material and connection form of the connection interface, thereby reducing the propagation of stress waves and achieving pyrotechnic impact isolation. However, most rigid impact isolation methods come at the cost of increasing the weight of the spacecraft.

[0003] Research shows that the cross-sectional shape and material properties of the material structure are the main factors affecting the propagation of stress waves, and the complexity of the honeycomb porous material structure is conducive to hindering the propagation of stress waves. By using the complexity of the honeycomb porous structure transmission path to construct the spacecraft-rocket interface, the stress wave impedance of the spacecraft-rocket interface can be effectively increased, the stress wave transmission can be reduced, and the pyrotechnic impact response can be isolated. At the same time, the area ratio is also a key parameter affecting the stress wave transmission rate, and the honeycomb structure has a large area ratio, which can effectively reduce the transmission of stress waves and facilitate the realization of impact isolation. In summary, for the pyrotechnic impact problem caused by the separation of the spacecraft and the rocket, the honeycomb porous material structure can be introduced to effectively suppress the impact response. SUMMARY

[0004] In order to solve the above technical problems, the present application provides a pyrotechnic impact isolation method using honeycomb porous material. The present application uses the honeycomb porous material structure for spacecraft impact isolation. By adding a honeycomb porous structure, when the spacecraft separates from the rocket, the complexity of the honeycomb porous structure transmission path is used to construct the spacecraft-rocket interface, the stress wave impedance of the spacecraft-rocket interface is increased, the stress wave transmission is reduced, and the pyrotechnic impact response is reduced. The spacecraft can be effectively protected to achieve effective isolation and suppression of pyrotechnic impact, improve the traditional spacecraft-rocket connection method, obtain ideal impact isolation effect with small weight increase, and ensure certain connection stiffness and reduce launch cost.

[0005] To solve the above technical problems, the technical scheme adopted by the present application is as follows:

[0006] A pyrotechnic impact isolation method using cellular porous material, comprising the following steps:

[0007] S1: According to the component level impact simulation experiment of satellite-rocket separation, the satellite component impact isolation requirement is obtained, and then the impact response size of the non-cushion structure and the allowable impact response size are obtained;

[0008] S2: A cellular porous structure is added at the bottom end of the spacecraft base to obtain an improved satellite-rocket interface;

[0009] S3: According to the requirements of the space satellite launcher, a positive triangular wave excitation is added at the explosion bolt, the satellite-rocket interface assembly is discretized and dynamically modeled, and the simulated explosion separation impact load is applied according to the actual launch requirements. The transient dynamics analysis is carried out by using the model to obtain the acceleration transient response of the satellite-rocket interface;

[0010] S4: The obtained model is analyzed by finite element method, the acceleration response of the satellite and the rocket end is extracted, the impact response spectrum attenuation is analyzed, whether the improved geometric model meets the impact requirements is verified, if it meets the requirements, the improved geometric model is obtained, and the design is completed; if it does not meet the requirements, the model parameters are revised and improved, and the above steps are repeated until the impact requirements are met.

[0011] Preferably, the geometric size of the cellular porous structure matches the size of the satellite-rocket interface.

[0012] Preferably, the material of the cellular porous structure is structural steel or aluminum alloy.

[0013] Compared with the prior art, the present application has the beneficial effects that:

[0014] The present application uses the cellular porous material structure for aircraft impact isolation. By adding a cellular porous structure, when the spacecraft separates from the rocket, the satellite-rocket interface is constructed by using the complexity of the transmission path of the cellular porous structure, the stress wave impedance of the satellite-rocket interface is increased, the stress wave transmission is reduced, and the pyrotechnic impact response is reduced. The spacecraft can be effectively protected to effectively isolate and suppress the pyrotechnic impact, improve the traditional satellite-rocket connection method, obtain the ideal impact isolation effect with very small weight increase, and also ensure a certain connection stiffness and reduce the launch cost. BRIEF DESCRIPTION OF DRAWINGS

[0015] Figure 1 is a system flow chart of an embodiment of the present application;

[0016] Figure 2 is an assembly drawing of the satellite-rocket interface and the cellular porous material model;

[0017] Figure 3 is a geometric model diagram of the original structure of the satellite-rocket interface;

[0018] Figure 4 is a geometric model diagram of the satellite-rocket interface and the honeycomb porous material;

[0019] Figure 5 is a stress nephogram of the original structure of the satellite-rocket interface;

[0020] Figure 6 is a stress nephogram of the satellite-rocket interface with the improved honeycomb porous material;

[0021] Figure 7 is a diagram of the upper and lower bottom sampling points of the original structure of the satellite-rocket interface in finite element simulation;

[0022] Figure 8 is a diagram of the upper and lower bottom sampling points of the satellite-rocket interface and the honeycomb porous material in finite element simulation;

[0023] Figure 9 is an impact response spectrum diagram of the original structure of the satellite-rocket interface;

[0024] Figure 10 is an impact response spectrum diagram of the satellite-rocket interface and the honeycomb porous material.

[0025] In the figure: 1, spacecraft base; 2, explosive bolt hole; 3, honeycomb porous structure; 4, explosive bolt. DETAILED DESCRIPTION

[0026] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor fall within the scope of protection of the present application.

[0027] As a pyrotechnic impact isolation method using a honeycomb porous material, the method comprises the following steps:

[0028] S1: According to the component-level impact simulation test of satellite-rocket separation, the impact isolation requirement of the satellite component is obtained, and then the impact response size of the non-cushion structure and the allowable impact response size are obtained;

[0029] S2: As shown in Figure 2 , a honeycomb porous structure 3 is added to the bottom end of the spacecraft base 1, as shown in Figure 4As shown, a honeycomb porous structure 3 of specific geometric dimensions is added to the bottom of the spacecraft base 1 to improve the traditional star-rocket connection method. The complexity of the transmission path of the honeycomb porous structure 3 is used to construct the star-rocket interface, increasing the stress wave impedance at the star-rocket interface, reducing stress wave transmission, and lowering the pyrotechnic shock response, thereby achieving pyrotechnic shock response isolation. The length and width dimensions of this structure must match the dimensions of the star-rocket interface; the wall thickness, side length, cell size, and other dimensions can be used to adjust the magnitude of the shock response; its material is structural steel or aluminum alloy. Figure 3 , Figure 4 A side view comparison of the star-rocket connection interface before and after improvement reveals that a honeycomb porous structure of a certain thickness has been incorporated into the interface.

[0030] S3: respectively for Figure 3 The original interface structure shown and Figure 4 The improved star-rocket interface structure shown was discretized using finite element method and dynamic modeled to obtain... Figure 5 and Figure 6 The stress cloud diagram of the corresponding finite element model is shown and used for transient dynamics calculation. Figure 7 These are the sampling points on the top and bottom surfaces of the original structure of the star-rocket interface in the finite element simulation. Figure 8 To improve the sampling points on the upper and lower bottom surfaces of the honeycomb structure in finite element simulation of the rear star-rocket interface;

[0031] S3: According to Saint Vincent's principle, a set of uniformly distributed excitations are added at the explosive bolt 4, and a simulated explosive separation impact load is applied according to the actual launch requirements to simulate the impact excitation generated by the explosion of the explosive bolt 4 when the satellite separates from the rocket. The transient dynamic analysis is performed using this model to obtain the acceleration transient response of the satellite-rocket interface.

[0032] The separation impact load applied in step S3 is a triangular wave excitation, and the magnitude and pulse width of the excitation are determined by the relevant parameters of the actual pyrotechnic device.

[0033] S4: Extract Figure 7 , Figure 8 The transient acceleration responses of the satellite and rocket at the location indicated by the middle arrow were analyzed and processed to obtain the impact response spectra at both sides of the two models, as shown below. Figure 9 and Figure 10 As shown in the figure. Simulation results show that in the original structure of the spacecraft-rocket interface, the relative coefficients Er and Mr in the impact response spectrum decrease as the stress wave generated by the pyrotechnic impact propagates from the explosion side of the lower bottom surface to the non-explosion side of the upper bottom surface. The average interface attenuation from the lower bottom surface to the upper bottom surface is 23.00%, and the maximum value is 16.49%. In the spacecraft with a honeycomb porous material structure, the average interface attenuation from the lower honeycomb bottom surface to the upper bottom surface of the base is 33.62%, and the maximum value is 42.12%, which meets the impact isolation requirements of actual launch missions.

[0034] In order to obtain the quantified shock response spectrum attenuation rate, two dimensionless coefficients not considering frequency are introduced:

[0035]

[0036] In the formula, Er is the average relative coefficient of the shock response spectrum in the whole frequency range, Mr is the maximum relative coefficient of the shock response spectrum, and SRSb(f) is the base SRS. Here, the shock response spectrum of the original structure is defined as the base shock response spectrum.

[0037] The simulation results show that in the original structure of the satellite-rocket interface, the relative coefficients Er and Mr in the shock response spectrum are attenuated with the transmission of the stress wave generated by the pyrotechnic shock from the lower bottom side of the explosion to the upper bottom side of the non-explosion. The average attenuation value from the lower bottom to the upper bottom is 23.00%, and the maximum value is 16.94%. In the spacecraft with the added honeycomb porous material structure, the average attenuation value from the lower bottom of the honeycomb to the upper bottom of the base is 33.62%, and the maximum value is 42.12%, which has reached the requirement of the shock isolation for the actual launch mission. The results show that after the addition of the honeycomb porous structure 3 at the satellite-rocket connection interface, the stress wave has been obviously reflected, and the shock response has been fully suppressed. The results prove that the pyrotechnic shock isolation device based on the honeycomb porous material can effectively realize the pyrotechnic shock isolation.

[0038] The above only describes the preferred embodiments of the present application in detail, but the present application is not limited to the above embodiments, and various changes can be made within the knowledge of those skilled in the art without departing from the purpose of the present application, and all the changes shall be included in the protection scope of the present application.

Claims

1. A method for shock isolation using cellular porous materials, characterized by: The method comprises the following steps: S1: obtaining the impact isolation requirement of the satellite components according to the component level impact simulation experiment of the satellite-rocket separation, and then obtaining the impact response size of the non-cushion structure and the allowable impact response size; S2: adding a honeycomb porous structure (3) at the bottom end of the spacecraft base (1) to obtain an improved satellite-rocket interface; S3: according to the requirements of the space satellite launcher, a positive triangular wave excitation is added at the explosive bolt (4), the satellite-rocket interface of the component is discretized and dynamically modeled by using the finite element method, and the simulated explosive separation impact load is applied according to the actual launch requirements, the transient dynamic analysis is carried out by using the model to obtain the acceleration transient response of the satellite-rocket interface; S4: performing finite element analysis on the obtained model, extracting the acceleration response of the satellite and the rocket end, analyzing the impact response spectrum attenuation, verifying whether the improved geometric model meets the impact requirements, if yes, the improved geometric model is obtained, and the design is completed; if not, the model parameters are revised and improved, and the above steps are repeated until the impact requirements are met.

2. A method for shock isolation of a pyrotechnic device using cellular porous material according to claim 1, characterized in that: The geometric size of the honeycomb porous structure (3) matches the size of the satellite-rocket interface.

3. The method of claim 1, wherein the cellular porous material is selected from the group consisting of: a cellular metal, a cellular ceramic, a cellular polymer, a cellular composite, and combinations thereof. The honeycomb porous structure (3) is made of structural steel or aluminum alloy.

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