Spacecraft pyrotechnic impact response prediction method based on impact load coupling solution

Abstract

The invention discloses a spacecraft pyrotechnic impact response prediction method based on impact load coupling solution, and the method comprises the steps: building a combustion gas flow field calculation model formed by gunpowder combustion based on a classical inner trajectory theory according to a filling structure of a pyrotechnic impact source, and solving to obtain a time domain impact load formed by energetic material combustion; secondly, establishing a finite element model, taking the time-domain impact load as the input of the finite element model, solving the finite element model to obtain the change of a system structure, and obtaining the force function and the pyrotechnic impact response of each part of the spacecraft; and finally, based on a load transfer method, establishing a coupling calculation framework of the gas flow field and the system structure field, taking the time domain impact load obtained by solving the flow field as the input of a finite element model at the beginning of each time step, and feeding back the system structure change obtained by the finite element model to a flow field calculation model at the end of each time step. Coupling calculation and real-time updating of the time domain impact load are achieved, and a flow field-structure field coupling calculation framework is closed.

Description

Technical Field

[0001] The invention belongs to the technical field of pyrotechnic impact technology, and in particular relates to a method for predicting the pyrotechnic impact response of a spacecraft based on the coupled solution of impact loads. Background Art

[0002] The pyrotechnic separation device is a pyrotechnic product. It drives the movement of functional mechanisms through the energy released by the combustion of energetic materials (gunpowder or explosives, etc.), and realizes functions such as rocket stage separation, satellite-rocket separation, payload release, and accessory deployment. It is widely used in spacecraft. Due to the fast combustion speed of energetic materials and the strong movement response of functional mechanisms, the working process of the pyrotechnic separation device will produce transient, high-frequency, and high-volume pyrotechnic shocks, which will damage instruments and equipment containing shock-sensitive components and brittle materials, leading to mission failure or even catastrophic accidents. Therefore, it is necessary to numerically simulate the pyrotechnic shock process to provide an effective reference for spacecraft structural design and safety protection.

[0003] At present, the commonly used method is to establish an explicit finite element model to simulate and predict the pyrotechnic impact process. This method requires providing an impact load as the input of the model. There are three main load input methods: (1) Obtaining the load function through classical waveform inversion calculation, but the impact response obtained by this method has certain differences in frequency and magnitude compared with the real impact response. (2) Using test data (such as pressure, acceleration response, etc.) as impact load. This method requires a lot of ground test support, which is time-consuming, costly, and has poor repeatability. (3) Directly solving the pyrotechnic impact process to obtain the time domain impact load. However, due to the complex coupling between the flow field and the system structure during the pyrotechnic impact process, the accuracy of solving the flow field and obtaining the impact load only through the combustion / explosion theory of energetic materials needs to be improved, and the ability to capture the characteristic parameters of the pyrotechnic impact process is still insufficient. Summary of the invention

[0004] The present invention proposes a method for predicting the pyrotechnic shock response of a spacecraft based on the coupled solution of the shock load. By means of the coupled calculation between the gas flow field of the energetic material and the system structure field during the pyrotechnic shock process, the pyrotechnic shock load solution capability, the pyrotechnic shock response prediction capability and the comprehensive capture capability of the characteristic parameters are improved, thereby solving the problem of low accuracy in directly solving the shock load and weak comprehensive capture capability of the characteristic parameters caused by the complex coupling between the gas flow field and the system structure during the pyrotechnic shock process.

[0005] The technical solution to realize the present invention is: a method for predicting the pyrotechnic shock response of a spacecraft based on the coupled solution of the shock load, comprising the following steps:

[0006] Step S10, according to the loading structure of the pyrotechnic impact source and based on the classical interior ballistic theory, a calculation model of the gas flow field formed by the combustion of gunpowder is established; in the first time step of the calculation time domain, the initial parameters of the gas flow field calculation model are set, the characteristic parameters of the gas flow field are solved, and the initial solution of the time domain impact load of the pyrotechnic impact process is obtained.

[0007] Step S20, establish a finite element model including the pyrotechnic impact source loading structure and the spacecraft system structure, and set the time domain impact load as p 2,i ; The initial solution p of the time domain impact load for the first time step in the time domain will be calculated 2,1 As the input of the finite element model; set the initial parameters of the finite element model, solve the finite element model, and obtain the motion process of the moving body and the structural field changes of the pyrotechnic impact source and the spacecraft system at the end of the first time step.

[0008] Step S30, based on the load transfer method, a coupling calculation framework of the gas flow field calculation model and the finite element model is established. Starting from the first time step in the calculation time domain, the calculation results of the finite element model are fed back to the flow field calculation model at the end of each time step. At the beginning of the next time step, the coupling solution of the time domain impact load of this time step is solved based on the parameters obtained in the previous time step, and the coupling solution is used as the load input of the finite element model to realize the closed loop of the flow field-structure field coupling calculation framework.

[0009] Step S40: convert the time domain impact load p 2,i The coupled solution of is used as the input of the finite element model to obtain the force function F of each part of the spacecraft system in each time step. i , and predict the pyrotechnic shock response A of each part of the spacecraft system i , in order to measure the pyrotechnic impact strength and protection capability of spacecraft, and provide an effective reference for spacecraft structural design and safety protection.

[0010] Compared with the prior art, the present invention has the following significant advantages:

[0011] (1) Based on the coupled solution of impact load, the changes in the system structure field are fed back to the gas flow field through the finite element model calculation in each time step of the calculation time domain, and the coupled solution of the impact load in the time domain can be obtained. The coupled solution of the impact load in the time domain is more accurate and can better reflect the influence of factors such as the combustion of energetic materials, the movement of moving bodies, the friction between structures, and the nonlinear deformation of the buffer device on the impact load during the pyrotechnic impact process.

[0012] (2) Based on the coupled solution of impact loads, the pyrotechnic impact loads will be updated and solved in real time at each time step in the calculation time domain. The obtained coupled solution is input into the finite element model to obtain more accurate prediction results of system structural field changes and pyrotechnic impact response.

[0013] (3) The coupled calculation framework includes the gas flow field of energetic materials and the system structure field. Through coupled calculation, it can comprehensively capture the combustion state of gunpowder in the flow field, the gas pressure distribution, the motion of moving bodies in the gas flow and structure field, the friction between structures, the nonlinear deformation of the buffer device, and the conservation of energy conversion between the flow field and the structure field, thus significantly improving the ability to capture the characteristic parameters of the pyrotechnic impact process. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] Figure 1 The figure is a flow chart of a method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to an embodiment of the present invention.

[0015] Figure 2 It is a schematic diagram of a finite element model of an pyrotechnic impact source loading structure and a spacecraft system structure according to an embodiment of the present invention.

[0016] Figure 3 Schematic diagram of the acquisition position of the force function according to an embodiment of the present invention.

[0017] Figure 4 The figure is a schematic diagram of the acquisition position of the pyrotechnic impact response according to an embodiment of the present invention.

[0018] Figure 5 FIG. 4 is a curve showing the development of the pyrotechnic impact pressure load over time according to an embodiment of the present invention.

[0019] Figure 6 FIG. 4 is a curve showing the development of the pyrotechnic impact force function over time according to an embodiment of the present invention.

[0020] Figure 7 FIG. 4 is a curve showing the development of pyrotechnic impact response over time according to an embodiment of the present invention.

[0021] Description of reference numerals:

[0022] 1. High-pressure chamber shell; 2. Low-pressure chamber shell; 3. Piston; 4. Push rod; 5. Spacecraft system end; 6. Spray hole; 7. Buffer device; 8. Shear diaphragm; 9. Separator. DETAILED DESCRIPTION

[0023] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0024] It should be noted that all directional indications in the embodiments of the present invention (such as up, down, left, right, front, back, etc.) are only used to explain the relative position relationship, movement status, etc. between the components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0025] In addition, in the present invention, the descriptions such as "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" or "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0026] In the present invention, unless otherwise clearly specified and limited, the terms "connection", "fixation" and the like should be understood in a broad sense. For example, "fixation" can be a fixed connection, a detachable connection, or an integral connection; "connection" can be a mechanical connection or an electrical connection. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0027] In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in the field can implement it. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0028] The following will further introduce the specific implementation method, as well as the technical difficulties and inventive points of this invention in combination with this design example.

[0029] Figure 1 A schematic flow chart of a method for predicting a spacecraft pyrotechnic shock response based on a coupled solution of shock loads according to an embodiment of the present invention is shown. In this embodiment, gunpowder is used as an energetic material. The method for predicting a spacecraft pyrotechnic shock response based on a coupled solution of shock loads comprises the following steps:

[0030] Step S10, according to the loading structure of the pyrotechnic impact source and based on the classical interior ballistic theory, a calculation model of the gas flow field formed by the combustion of gunpowder is established; in the first time step of the calculation time domain, the initial parameters of the gas flow field calculation model are set, the characteristic parameters of the gas flow field are solved, and the initial solution of the time domain impact load of the pyrotechnic impact process is obtained.

[0031] Specifically, a programming method can be used to establish and solve the calculation model of the gas flow field formed by the combustion of gunpowder, and the programming platform can be a subroutine interface of the commercial software ABAQUS.

[0032] according to Figure 2 The loading structure of the pyrotechnic impact source shown in the figure, in this embodiment, the impact load is expressed in the form of the gunpowder gas pressure inside the low-pressure chamber shell, and the calculation formula is shown in formula (1):

[0033]

[0034] Where i is the time step number, i=1 represents the first time step in the calculation domain; p 2,i is the pressure of the gunpowder gas inside the low-pressure chamber shell at the i-th time step, that is, the time domain impact load; f is the gunpowder force; ω is the gunpowder mass; η i V is the relative gas flow rate from the high-pressure chamber shell to the low-pressure chamber shell at the i-th time step; 2,i is the volume of the low-pressure chamber shell at the i-th time step. When i=1, V 2,1 Represents the initial volume of the low pressure chamber shell.

[0035] Step S20, establish a finite element model including the pyrotechnic impact source loading structure and the spacecraft system structure, and set the time domain impact load as p 2,i ; The initial solution p of the time domain impact load for the first time step in the time domain will be calculated 2,1 As the input of the finite element model; set the initial parameters of the finite element model, solve the finite element model, and obtain the motion process of the moving body and the structural field changes of the pyrotechnic impact source and the spacecraft system at the end of the first time step.

[0036] In this embodiment, the finite element model is as follows: Figure 2 As shown, the finite element model includes a high-pressure chamber shell 1, a low-pressure chamber shell 2, a piston 3, a push rod 4, a spacecraft system end 5 and a separator 9. Explosive powder is loaded in the high-pressure chamber shell 1, the high-pressure chamber shell 1 and the low-pressure chamber shell 2 are fixedly connected, and a spray hole 6 is provided at the center of the connecting end surface, and the high-pressure chamber shell 1 and the low-pressure chamber shell 2 are connected through the spray hole 6. The high-pressure chamber shell 1 and the low-pressure chamber shell 2 constitute the pyrotechnic impact source. The push rod 4 with the piston 3 fixedly connected at the end is located in the low-pressure chamber shell 2, the piston 3 is on the same side as the high-pressure chamber shell 1, and a through hole is opened in the center of the other end surface of the low-pressure chamber shell 2 as an outlet. The push rod 4 extends from the outlet to push the separator 9 outside the low-pressure chamber shell 2, and the spacecraft system end 5 is located behind the separator 9.

[0037] During the operation of the spacecraft, the gunpowder loaded in the high-pressure chamber housing 1 is ignited and starts to burn, generating high-temperature and high-pressure combustion gas. The combustion gas flows into the low-pressure chamber housing 2 through the nozzle 6, and acts on the piston 3 in the form of pressure, pushing the piston 3 and the push rod 4 to move backward, and pushing the separation member 9 to separate. This process is accompanied by the impact of the piston 3 on the low-pressure chamber housing 2 and the impact of the separation member 9 on the spacecraft system end 5.

[0038] Under certain working conditions, a shear diaphragm 8 is further provided at the outlet of the impacted end of the low-pressure chamber housing 2 to control the starting pressure of the push rod 4 .

[0039] Under certain working conditions, a buffer device 7 is further provided at the impacted end of the low-pressure chamber housing 2 to buffer the pyrotechnic impact from the piston 3 .

[0040] A finite element model including the pyrotechnic impact source loading structure and the spacecraft system structure is established. Specifically, commercial software ABAQUS can be used, and the programming platform should be ensured to match the software used in the finite element model.

[0041] According to the calculation results of the finite element model, at the end of the first time step, the displacement l1, velocity v1, acceleration a1 of the moving body are extracted, and the kinetic energy E of the moving body is extracted. K,1 , Friction energy consumption of moving body E F,1 , Energy absorption of buffer device E ED,1 .

[0042] Step S30, based on the load transfer method, a coupling calculation framework of the gas flow field calculation model and the finite element model is established. Starting from the first time step in the calculation time domain, the calculation results of the finite element model are fed back to the flow field calculation model at the end of each time step. At the beginning of the next time step, the coupling solution of the time domain impact load of this time step is solved based on the parameters obtained in the previous time step, and the coupling solution is used as the load input of the finite element model to realize the closed loop of the flow field-structure field coupling calculation framework.

[0043] The coupling calculation framework considers the coupling effect of flow field and structural field in each time step i. The time domain impact load p expressed in the form of pressure 2,i The coupling solution is the real impact load borne by the spacecraft during the pyrotechnic impact process. According to different coupling strategies, the coupling framework can be divided into energy conservation coupling and dynamic coupling. Both calculation methods can realize the time domain impact load p 2,i Solving the coupled solution.

[0044] The energy conservation coupling framework calculation method is shown in equations (2), (3) and (4):

[0045] η i =f(p 2，i ), i = 1, 2, 3...n (2)

[0046]

[0047] p 2,i+1 =f(η i , E K,i , E F，i , E ED,i), i = 1, 2, 3...n (4)

[0048] Among them, E K,i is the kinetic energy of the moving body; E F,i E is the friction energy consumed by the moving body; ED,i is the energy absorption of the buffer device; η i is the relative gas flow rate flowing from the high-pressure chamber shell into the low-pressure chamber shell at the i-th time step; f represents the function; n represents the total number of time steps.

[0049] The calculation formulas of the dynamic coupling framework are shown in equations (2), (5) and (6):

[0050]

[0051] p 2，i+1 =f(η i , l i , v i , a i ), i = 1, 2, 3...n (6)

[0052] Among them, l i is the displacement of the moving body; v i is the speed of the moving body; a i is the acceleration of the moving body.

[0053] Step S40: convert the time domain impact load p 2,i The coupled solution of is used as the input of the finite element model to obtain the force function F of each part of the spacecraft system in each time step. i , and predict the pyrotechnic shock response A of each part of the spacecraft system i , in order to measure the pyrotechnic impact strength and protection capability of spacecraft, and provide an effective reference for spacecraft structural design and safety protection.

[0054] The calculation formulas are shown in formula (7) and formula (8):

[0055] F i =f(p 2，i ), i = 1, 2, 3...n (7)

[0056] A i =f(F i ), i = 1, 2, 3...n (8)

[0057] It is worth noting that the spacecraft system can be a whole spacecraft that can independently undertake space missions, such as the entire rocket, a satellite system separated from the rocket, etc., or an independent pyrotechnic separation device, such as an explosive nut, a separation nut, an unlocking push rod, etc. In this embodiment, the spacecraft system is an unlocking push rod using gunpowder as the energetic material, but it should not be limited to this and can be adjusted according to actual conditions.

[0058] In this embodiment, the force function of each part of the spacecraft system during the pyrotechnic impact process can be obtained through the coupling calculation framework of the gas flow field calculation model and the finite element model. For example, the force function F′ of the impacted end of the low-pressure chamber shell 2 can be obtained. i , the force function F″ of the spacecraft system end 5 can be obtained i ,like Figure 3 shown.

[0059] In this embodiment, the pyrotechnic shock response of each part of the spacecraft system during the pyrotechnic shock process can be obtained through the coupling calculation framework of the gas flow field calculation model and the finite element model. For example, the pyrotechnic shock response A′ of the impacted end of the low-pressure chamber shell 2 can be obtained. i , the pyrotechnic shock response A″ of the spacecraft system end 5 can be obtained i ,like Figure 4 shown.

[0060] In this embodiment, the curve of the development of the pyrotechnic impact load expressed in the form of pressure over time is as follows: Figure 5 As shown. The curve in the figure can be divided into three stages. Before about 5.65ms, the impact load showed a trend of first rising and then falling. This stage is the most important stage of pyrotechnic impact. The gunpowder continues to burn in the high-pressure chamber shell 1 and the low-pressure chamber shell 2. The gas in the low-pressure chamber shell 2 pushes the piston 3 and the push rod 4 to move, the combustion space V2,i expands, the separator 9 separates, and the buffer device 7 is compressed and deformed by the impact. Until 5.65ms, the movement of the piston 3 and the push rod 4 stops under the buffering action of the buffer device 7. At this stage, part of the energy contained in the gunpowder gas is converted into the kinetic energy E of the moving body. K,i , friction energy consumption of moving body E F,i , Energy absorption of the buffer device E ED，i. As shown in formula (4), these characteristic parameters will be fed back to the impact load solution in the coupling framework to improve the accuracy of the calculation results. From 5.65ms to 7.38ms, the impact load enters the second stage. In this stage, the movement of the piston 3 and the push rod 4 stops, the deformation of the buffer device 7 enters a stable state, and the feedback of the system structure field on the gas flow field can be almost ignored. However, the gunpowder in the high-pressure chamber shell 1 and the low-pressure chamber shell 2 is not completely burned until 7.38ms, causing the impact load to rise again, forming a second impact on the system. After 7.38ms, the impact load enters the third stage. Since the finite element simulation does not consider gas leakage, when the entire system is in a stable state, the impact load remains constant.

[0061] In this embodiment, a buffer device 7 is provided at the impacted end of the low-pressure chamber housing 2 to buffer the pyrotechnic impact. The impact force function F a And the impact force function F borne by the impacted end of the low-pressure chamber shell 2 b like Figure 6 As shown in the figure, the buffer device 7 bears the impact force before the impacted end of the low-pressure chamber shell 2, and absorbs the kinetic energy E carried by the movement of the piston 3 and the push rod 4 through its own plastic telescopic deformation. K，i When the impact load is transmitted to the impacted end of the low-pressure chamber housing 2, the impact force is buffered and weakened by the buffer device 7, showing a trend of multiple peaks and troughs corresponding to the plastic telescopic deformation process. Until the buffering capacity of the buffer device 7 reaches the upper limit, the impact force borne by the impacted end of the low-pressure chamber housing 2 increases sharply, and then gradually decreases.

[0062] In this embodiment, the pyrotechnic impact response of the impacted end of the low-pressure chamber housing 2 is as follows: Figure 7 As shown in the figure, it can be seen that before 5.65ms, the buffer device 7 effectively exerted its buffering capacity, and the pyrotechnic shock response of the impacted end of the low-pressure chamber shell 2 was in a lower range. After the buffering capacity of the buffer device 7 reached the upper limit, the impact force of the impacted end of the low-pressure chamber shell 2 increased sharply, and the pyrotechnic shock response also experienced a sharp fluctuation. This is also the moment when the shock response is the highest in the entire pyrotechnic shock process.

[0063] The above are only specific embodiments of the present invention, but the protection scope of the present invention is not limited thereto. The protection scope of the present invention shall be based on the protection scope of the claims.

Claims

1. A method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution, characterized in that: The following steps are involved: Step S10, according to the loading structure of the pyrotechnic impact source and based on the classical interior ballistic theory, a calculation model of the gas flow field formed by the combustion of gunpowder is established; in the first time step of the calculation time domain, the initial parameters of the gas flow field calculation model are set, the characteristic parameters of the gas flow field are solved, and the initial solution of the time domain impact load of the pyrotechnic impact process is obtained; Step S20, establish a finite element model including the pyrotechnic impact source loading structure and the spacecraft system structure, and set the time domain impact load as p 2,i ; The initial solution p of the time domain impact load for the first time step in the time domain will be calculated 2,1 As the input of the finite element model; set the initial parameters of the finite element model, solve the finite element model, and obtain the motion process of the moving body and the pyrotechnic impact source and the structural field changes of the spacecraft system at the end of the first time step; Step S30, based on the load transfer method, a coupling calculation framework of the gas flow field calculation model and the finite element model is established, starting from the first time step in the calculation time domain, at the end of each time step, the calculation result of the finite element model is fed back to the flow field calculation model, and at the beginning of the next time step, the coupling solution of the time domain impact load of this time step is solved based on the parameters obtained in the previous time step, and the coupling solution is used as the load input of the finite element model to realize the closed loop of the flow field-structure field coupling calculation framework; Step S40: convert the time domain impact load p 2,i The coupled solution of is used as the input of the finite element model to obtain the force function F of each part of the spacecraft system in each time step. i , and predict the pyrotechnic shock response Ai of each part of the spacecraft system to measure the pyrotechnic shock strength and protection capability of the spacecraft, providing an effective reference for the spacecraft structure design and safety protection.

2. The method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to claim 1 is characterized in that: In step S10, a calculation model of the gas flow field formed by the combustion of gunpowder is established, which is specifically as follows: The impact load is expressed in the form of the gunpowder gas pressure inside the low-pressure chamber shell: Where i is the time step number, i=1 represents the first time step in the calculation domain; p 2,i is the pressure of the gunpowder gas inside the low-pressure chamber shell at the i-th time step, which is the time domain impact load; f is the gunpowder force; ω is the gunpowder mass; η i V is the relative gas flow rate from the high-pressure chamber shell to the low-pressure chamber shell at the i-th time step; 2,i is the volume of the low-pressure chamber shell at the i-th time step.

3. The method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to claim 2 is characterized in that: In step S20, a finite element model including the pyrotechnic impact source loading structure and the spacecraft system structure is established, as follows: The finite element model comprises a high-pressure chamber shell (1), a low-pressure chamber shell (2), a piston (3), a push rod (4), a spacecraft system end (5) and a separator (9); explosive powder is loaded in the high-pressure chamber shell (1); the high-pressure chamber shell (1) and the low-pressure chamber shell (2) are fixedly connected; a spray hole 6 is provided at the center of the connecting end surface; the high-pressure chamber shell (1) and the low-pressure chamber shell (2) are connected through the spray hole 6; the high-pressure chamber shell (1) and the low-pressure chamber shell (2) constitute an explosive impact source; the push rod (4) with the piston (3) fixedly connected at the end is located in the low-pressure chamber shell (2); the piston (3) and the high-pressure chamber shell (1) are on the same side; a through hole is provided at the center of the end surface on the other side of the low-pressure chamber shell (2) as an outlet; the push rod (4) extends from the outlet to push the separator (9) outside the low-pressure chamber shell (2); and the spacecraft system end (5) is located behind the separator (9); During the operation of the spacecraft, the gunpowder loaded in the high-pressure chamber shell (1) is ignited and starts to burn, generating high-temperature and high-pressure combustion gas; the combustion gas flows into the low-pressure chamber shell (2) through the nozzle hole 6, acts on the piston (3) in the form of pressure, pushes the piston (3) and the push rod (4) to move backward, and pushes the separation member (9) to separate; this process is accompanied by the impact of the piston (3) on the low-pressure chamber shell (2) and the impact of the separation member (9) on the spacecraft system end (5); According to the calculation results of the finite element model, at the end of the first time step, the displacement l1, velocity v1, acceleration a1 of the moving body are extracted, and the kinetic energy E of the moving body is extracted. K,1 , Friction energy consumption of moving body E F,1 , Energy absorption of buffer device E ED,1 .

4. The method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to claim 3 is characterized in that: The finite element model also includes a shear diaphragm (8) arranged at the outlet of the impacted end of the low-pressure chamber housing (2) for controlling the starting pressure of the push rod (4).

5. The method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to claim 4 is characterized in that: The finite element model also includes a buffer device (7) arranged at the outlet of the impact-receiving end of the low-pressure chamber housing (2) for buffering the pyrotechnic impact from the piston (3).

6. The method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to claim 5 is characterized in that: In step S30, based on the load transfer method, a coupling calculation framework of the gas flow field calculation model and the finite element model is established. Starting from the first time step in the calculation time domain, the calculation results of the finite element model are fed back to the flow field calculation model at the end of each time step. At the beginning of the next time step, the coupling solution of the time domain impact load of this time step is solved based on the parameters obtained in the previous time step, and the coupling solution is used as the load input of the finite element model to realize the closed loop of the flow field-structure field coupling calculation framework, as follows: The coupled calculation framework considers the coupling effect of flow field and structural field in each time step i. The time domain impact load p expressed in the form of pressure 2,i The coupled solution of is the real impact load borne by the spacecraft during the pyrotechnic impact process; According to different coupling strategies, the coupling framework can be divided into energy conservation coupling and dynamic coupling. Both calculation methods can realize the solution of the coupling solution of the time domain impact load p2,i; The energy conservation coupling framework calculation method is shown in equations (2), (3) and (4): or i =f(p 2，i ),i=1,2,3...n (2) p 2,i+1 =f(η i ,AND K,i ,AND F，i ,AND ED,i ),i=1,2,3...n (4) Among them, E K,i is the kinetic energy of the moving body; E F,i E is the friction energy consumed by the moving body; ED,i is the energy absorption of the buffer device; η i is the relative gas flow rate from the high-pressure chamber shell to the low-pressure chamber shell at the i-th time step; f represents the function; n represents the total number of time steps; The calculation formulas of the dynamic coupling framework are shown in equations (2), (5) and (6): p 2，i+1 =f(η i ,l i ,v i ,a i ),i=1,2,3...n (6) Among them, l i is the displacement of the moving body; v i is the speed of the moving body; a i is the acceleration of the moving body.

7. The method for predicting spacecraft pyrotechnic shock response based on shock load coupling solution according to claim 6 is characterized in that: In step S40, the time domain impact load p 2，i The coupled solution of is used as the input of the finite element model to obtain the force function F of each part of the spacecraft system in each time step. i , and predict the pyrotechnic shock response A of each part of the spacecraft system i , as follows: The calculation formulas are shown in formula (7) and formula (8): F i =f(p 2，i ),i=1,2,3...n (7) A i =f(F i ),i=1,2,3...n (8)。

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