Simulation Method of Flow Field of Rocket Gas Jet Impacting Vertical Launching Device

Abstract

This invention discloses a method for simulating the flow field of rocket exhaust jet impacting a vertical launch device, belonging to rocket launch technology. This invention establishes a flow model of the rocket's takeoff force and thermal environment based on nested mesh technology. The inviscid flux of the multi-component Navier-Stokes equations containing exhaust gas transport is discretized using an improved Roe scheme, and the inviscid flux at the grid cell boundaries is obtained. After setting boundary parameters for the flow model, a Realizable k-ε turbulence model is used to solve the viscous dissipation in the Navier-Stokes equations, and the viscous flux is discretized using a central difference method to complete the entire calculation. The advantages of this invention compared to existing technologies are reduced computation time; the use of nested meshes, the improved Roe scheme, and pressure ramp techniques allows for better simulation of actual conditions, effectively solving the problem of simulating the exhaust jet flow field during ignition at different attitude angles, altitudes, and velocities of ejected rockets.

Description

TECHNICAL FIELD

[0001] The present application relates to a rocket launching technology, in particular to a method for simulating the flow field of a rocket gas jet impacting a vertical launching device. BACKGROUND

[0002] When a rocket is launched, high-temperature and high-speed gas jets are discharged from the engine combustion chamber, which will have a strong impact on the walls of the launching vehicle, launching cylinder and instrument cabin, and will have a strong impact on the walls of the launching vehicle, launching cylinder and instrument cabin. The effect of dynamic and thermal effect, especially when the rocket is disturbed to produce a large attitude angle, the oblique impact of the gas flow is easy to induce the violent vibration of the launching device. Therefore, in the overall design, the problems of ignition height, lightweight of launching device, stability and safety must be solved. At present, the equipment using vertical cold launching mode at home and abroad includes Russian S300 and S400 air defense missile system, birch, salmat ballistic missile system. American sea sparrow air defense missile system, missile 3, and peace guardian ballistic missile. French "purple" air defense missile system. China's red flag series air defense missile system, Dongfeng series ballistic missile system all use vertical cold launching mode. Compared with hot launching, after the cold launching of the vehicle-mounted missile, due to the disturbance of the cylinder port, the attitude angle of the missile body changes, for example, the missile body produces a certain deflection angle with the vertical axis, and at the same time, it may also rotate. When the engine is ignited at a certain height from the launching cylinder, the high-temperature and high-speed gas jet will have a strong impact and thermal ablation effect on the ground launching device. Therefore, accurately predicting the dynamic and thermal effects caused by the impact of rocket gas jet on the launching cylinder and instrument cabin has important reference significance for the stability design and heat protection design of the ground launching device, and is also conducive to improving the performance of the launching device.

[0003] Due to the harsh force-thermal environment of rocket launch, the vehicle protection material not only needs to withstand high temperature, but also needs to be lightweight to achieve the performance indicators of rapid mobility. If ground flight tests are carried out, there are many influencing factors of the vehicle vertical launch force-thermal environment, such as the attitude angle at ignition, the ignition height, etc. It can be seen that the cost of completing the experimental simulation of multiple states is high, and the cycle is long. In recent years, with the rapid development of computational fluid dynamics and the substantial improvement of computer performance, large-scale precise numerical simulation of the rocket launch gas jet impact effect has become an important development trend. In 2020, when a certain rocket in China was simulated by static numerical calculation, the error between the simulation data and the experimental data was about 10% under the condition of using 45 CPU core parallel operation for 2 months. Especially for large-caliber rocket missiles, the gas jet impact load is very serious during launch. At present, the following problems are mainly encountered when simulating the gas jet force-thermal environment: (1) The overall calculation flow region of the rocket launch device is very large, with a height of more than 60M and a grid size of 15 million structured grid elements, and the amount of operation is huge. (2) The rocket gas jet belongs to a highly under-expanded jet, and the flow field has jet shock wave, Mach disk normal shock wave, three-pronged shock wave, expansion wave and jet boundary contact discontinuity. Such a complex wave system structure is easy to cause numerical divergence. (3) The problem of moving boundary caused by the high-temperature and high-speed gas jet discharged during the rocket movement cannot be solved by the layer domain dynamic grid, and the non-structured grid reconstruction cannot control the grid size, scale and mass. SUMMARY

[0004] The purpose of the present application is to provide a rocket gas jet impact vertical launch device flow field simulation method to solve the impact launch cylinder and vehicle load and heat protection problems caused by the unsteady complex flow field changes of rocket takeoff.

[0005] The technical method to achieve the purpose of the present application is:

[0006] A rocket gas jet impact vertical launch device flow field simulation method, comprising the following steps:

[0007] Step 1, establish a three-dimensional model of the vehicle vertical launch device, the three-dimensional model comprising a vehicle body 1, a launch cylinder 2 and a rocket 3, and proceed to step 2.

[0008] Step 2, divide the established three-dimensional model of the vehicle vertical launch device into flow regions, and construct a moving nested grid system:

[0009] Divide the established three-dimensional model of the vehicle vertical launch device into flow regions, adopt a block grid generation strategy, first divide the entire flow region into multiple sub-domains, then generate calculation structured grids for each sub-domain, thereby constructing a moving nested grid system, and proceed to step 3.

[0010] Step 3, select a theoretical model according to the moving nested grid system and establish a flow model of the rocket take-off force and heat environment:

[0011] The multi-component N-S equation containing fuel gas transport is discretized into the improved Roe finite volume format, the positive factor value is determined, the inviscid flux at the cell boundary of the calculation structure grid is obtained by numerical integration-differentiation of the flow equation of the improved Roe finite volume format, and step 4 is entered.

[0012] Step 4, setting boundary parameters for the flow model of the rocket take-off force and heat environment:

[0013] The background grid and the moving grid of the moving nested grid system are defined, and the background grid and the moving grid are cut and holed respectively to construct the flow model of the moving boundary, the flow model of the rocket take-off force and heat environment and the rocket motion equation (known part of parameters) are solved, the flow parameters of the rocket engine combustion chamber are defined (wherein the pressure increases with time), and the atmospheric environment parameters are defined according to the atmospheric environment during launching, and step 5 is entered.

[0014] Step 5, solving the above-mentioned flow model of the rocket take-off force and heat environment and its boundary parameters, the flow parameters of the rocket engine combustion chamber and the atmospheric environment parameters to obtain the flow field results: the Realizable k-ε turbulence model is used to solve the viscous dissipation in the N-S equation, the central difference is used to disperse the viscous flux, the whole N-S equation solution is completed, and the Mach number, temperature and pressure flow field atlas and the load change curve of the gas jet impacting the launching cylinder are obtained.

[0015] Compared with the prior art, the present application has the following advantages:

[0016] (1) The flow field simulation method of the rocket gas jet impacting the vertical launching device of the present application effectively solves the problems of grid quality control, numerical iteration stability and parallel computing efficiency caused by the large motion range of the rocket body, large grid scale and six-degree-of-freedom motion boundary.

[0017] (2) The flow field simulation method of the rocket gas jet impacting the vertical launching device of the present application can solve the complex unsteady gas jet impact effect calculation of the launching rocket at different attitude angles, ignition height and ignition time, so that the simulation results are more realistic and efficient.

[0018] (3) The improved second-order Roe upwind format is used in the present application, which adds a positive factor compared with the original format. Compared with other formats, the improved Roe format has the total variation diminishing (TVD) property, small numerical dissipation, not only has the advantages of the original Roe format, but also can better calculate the shock strong discontinuity non-physical oscillation solution, especially suitable for strong shock problems such as Mach disk in rocket jet.

[0019] (4) The rocket gas jet impact vertical launching device flow field simulation method of the application establishes a reasonable moving nested grid model, so that the grid division of the fluid domain is more convenient, and the time of large-scale iterative operation of the force and heat environment of the gas flow impact under the rocket motion state is effectively saved.

[0020] (5) The rocket gas jet impact vertical launching device flow field simulation method of the application adopts the Realizable k-ε model which is improved on the basis of the standard k-ε equation, describes the turbulent vortex in the turbulent equation, improves the calculation accuracy of the adverse pressure gradient of the gas jet, and can also flexibly handle the grid size near the wall surface. BRIEF DESCRIPTION OF DRAWINGS

[0021] Figure 1 The rocket gas jet impact vertical launching device flow field simulation method structure flow chart.

[0022] Figure 2 The three-dimensional model of the vehicle-mounted vertical launching device.

[0023] Figure 3 The multi-nested grid block structured processing diagram of the vehicle-mounted vertical launching device.

[0024] Figure 4 The rocket motion domain nested grid diagram of the vehicle-mounted vertical launching device.

[0025] Figure 5 The Mach number field cloud chart of the vehicle-mounted vertical launching device at three typical moments after the rocket main engine works.

[0026] Figure 6 The temperature field cloud chart of the vehicle-mounted vertical launching device at three typical moments after the rocket engine works.

[0027] Figure 7 The pressure field cloud chart of the vehicle-mounted vertical launching device at three typical moments after the rocket engine works.

[0028] Figure 8 The vertical force change curve diagram of the rocket take-off process non-steady gas jet impact launching cylinder.

[0029] Figure 9 The engine time-pressure change curve diagram of the rocket take-off process. DETAILED DESCRIPTION

[0030] In order to illustrate the technical scheme and technical purpose of the application, the application will be further introduced below in combination with the drawings and specific embodiments.

[0031] In combination with the drawings Figure 1The application discloses a rocket gas jet flow field simulation method for impacting a vertical launching device.

[0032] Step 1: establishing a three-dimensional model of the vehicle-mounted vertical launching device.

[0033] Combined Figure 2 The three-dimensional model comprises a vehicle body 1, a launching barrel 2 and a rocket 3, the vehicle body 1 comprises a cabin body, a vehicle head, a tire and a chassis: the three-dimensional model parameters are consistent with an actual engineering, and the three-dimensional model is established in a 1:1 equal ratio with the actual engineering launching vehicle and the rocket.

[0034] The three-dimensional model parameters comprise the following: the geometric size of the vehicle body 1, the diameter, the length and the throat radius parameter of the rocket 3, the inner diameter, the outer diameter and the length of the launching barrel 2.

[0035] Go to step 2.

[0036] Step 2: dividing a flow region of the established three-dimensional model of the vehicle-mounted vertical launching device and constructing a moving nested grid system.

[0037] The flow region of the established three-dimensional model of the vehicle-mounted vertical launching device is divided, a block grid generation strategy is adopted, the entire flow region is divided into a plurality of subdomains, and then a calculation structure grid is generated for each subdomain, so that the moving nested grid system is constructed.

[0038] Combined Figure 3 Firstly, the flow region of the established three-dimensional model of the vehicle-mounted vertical launching device is divided, the flow region (namely the overall calculation domain) is subjected to block grid processing through the block grid generation strategy, namely the overall calculation domain is divided into four subdomains, which are a rocket moving flow domain, a launching barrel and an upper flow domain, a vehicle body and a vehicle body surrounding jet flow aggregation flow domain and a remaining flow domain, and then a calculation structure grid is generated for each subdomain. The grid area of the rocket moving flow domain is a main body grid, the grid area of the remaining flow domain and the grid area of the launching barrel and the upper flow domain are background grids, and the main body grid is connected with the two background grids to solve the motion boundary problem of the rocket taking off, so that the moving nested grid system is constructed.

[0039] The launching barrel flow domain and the vehicle body and the vehicle body surrounding jet flow aggregation flow domain need to be subjected to structured grid encryption processing, so as to ensure the orthogonality and smoothness of the grid, and the grid in the fluid region far away from the vehicle-mounted vertical launching device is transitioned from dense to sparse.

[0040] Go to step 3.

[0041] Step 3: selecting a theoretical model according to the moving nested grid system and establishing a flow model of the rocket taking-off force-thermal environment.

[0042] The multi-component N-S equation of the gas containing fuel is discretized into the improved Roe finite volume format, the size of the positive factor is determined, and the inviscid flux at the cell boundary of the calculation structure grid is obtained by numerically integrating the flow equation of the differential element from the improved Roe finite volume format.

[0043] Since the composition of the exhaust gas of the rocket engine is complex, the flow model of the rocket take-off force thermal environment is a multi-component flow model of the gas and air. The inlet pressure of the rocket engine increases with time, and the movement of the rocket is described by the rocket movement equation.

[0044] S3.1, Before establishing the flow model of the rocket take-off force thermal environment, the following basic assumptions are made for the gas jet: 1) the gas jet satisfies the continuous medium; 2) the gas jet is a compressible pure gas phase medium; 3) no chemical reaction occurs inside the gas jet; 4) a multi-component mixed flow model of gas and air is used, and the multi-component satisfies the ideal gas state equation.

[0045] S3.2, The above assumptions ensure that the gas is a Newtonian viscous fluid, so the multi-component transport equation of the gas is established as

[0046]

[0047] where Y a is the mass fraction of the rocket gas component a, is the flux of the physical quantity, such as is the diffusion term of the rocket gas component, and is the flux of the diffusion of the rocket gas component, R a is the net generation rate of the rocket gas component a through chemical reaction, S a is the generation rate caused by the dispersion phase of the rocket gas from the self-defined source term, t is the flight time of the multi-nozzle rocket, p is the density of the rocket gas, v is the flow velocity of the rocket gas, and N is the number of rocket gas components.

[0048] where the flux of the diffusion of the component in equation (1) is

[0049]

[0050] In equation (2), D a,m is the diffusion coefficient of component a in the mixed medium, D T,a is the divergence of the gas flow element, and T is the temperature of the rocket gas.

[0051] 3.3, The above assumption that the gas is a compressible gas is made, and the conservation form of the compressible Navier-Stokes equation (N-S equation) is established in the rectangular coordinate system as

[0052]

[0053] Wherein, V is the velocity vector of the rocket flight, S is the gas flow cross-sectional area, Q is the rocket gas combustion heat, is the generalized source term of the multi-nozzle rocket, is the direction vector in the X, Y, Z direction of the rectangular coordinate system, E, F and G are the rocket gas flow flux vectors in the X, Y, Z direction of the rectangular coordinate system, E v , F v and G v are the rocket gas viscous flux vectors in the X, Y, Z direction of the rectangular coordinate system.

[0054] S3.4, the Roe format of the improved N-S equation is discretized, and the inviscid flux at the element boundary of the calculation structure grid is solved:

[0055] The direct solving process of N-S equation is complicated and consumes a lot of calculation time, especially the multi-component transport condition increases the difficulty of calculation, the improved Roe format is innovatively put forward. The improved Roe format adds a positive factor to the original Roe format, so that it not only has the strong ability of the original Roe format to distinguish shock wave, slip flow and other physical discontinuities, but also has good positive definite nature, will not produce red jade phenomenon when calculating high-speed flow, is easy to expand to other hyperbolic systems, can better handle the situation of low-speed propagation of shock wave in calculation, and has high calculation efficiency. This makes it not only able to solve the shock wave in the inviscid flow field with high resolution, but also can accurately process shear flow and other problems in viscous flow field.

[0056] The inviscid flux at the boundary of the calculation structure grid is expanded as follows by using the basic form of Roe format:

[0057]

[0058] ξ=|U| (5)

[0059]

[0060]

[0061]

[0062]

[0063] Wherein, is the numerical inviscid flux of Roe format, β is the positive factor innovatively added in this paper, which is mainly related to the temperature and density of the gas, and β is valued between 0-0.5 according to different physical properties of the gas. Δ(*) is the change of physical quantity *, ξ is the basic upwind dissipation term, δp u is the interface correction pressure driven by velocity gradient, δp pThe interface correction pressure for the pressure gradient, δU u The interface correction velocity for the velocity gradient, δU p The interface correction velocity for the pressure gradient, n x , n y , n z The components of the interface normal vector on the X, Y and Z axes respectively, c is the sound speed of the atmospheric environment, U is the interface normal velocity, p is the density of the rocket gas, u, b and w are the X-direction, Y-direction and Z-direction rocket gas flow velocities respectively, p is the pressure, B is the total energy, and H is the total enthalpy of the rocket gas.

[0064] The Roe format has the advantages of the AUSM+ format and the Van Leer format, and overcomes the disadvantages of the two formats.

[0065] The flow model of the rocket take-off force-thermal environment is shown in formulas (1)-(9).

[0066] Go to step 4.

[0067] Step 4, boundary parameters are set for the flow model of the rocket take-off force-thermal environment:

[0068] The background grid and the moving grid of the moving nested grid system are defined, and the background grid and the moving grid are respectively cut and hole-digged to construct the flow model of the moving boundary. The rocket engine nozzle inlet boundary defines the pressure-time variation equation of the combustion chamber. In the previous calculation method of the rocket gas jet, the rocket discharges the gas jet at a fixed height to impact the ground launching device, and the engine sprays the gas flow at a constant total pressure, so that the simulation result deviates greatly from the actual situation. The root cause is that the initial shock wave of the rocket gas flow at a constant total pressure is very strong, and the impact load generated is also very large, which deviates from the actual situation and cannot observe the entire process of the rocket launch. The vehicle-mounted vertical launching device in the present application causes the force-thermal environment simulation by impacting with unsteady gas jet, which is closer to the real rocket movement and engine ignition pressure building conditions, so the prediction result is naturally more accurate.

[0069] S4.1, set the nested cutting boundary and the background grid:

[0070] As shown in Figure 4 , when the rocket moves in the flow domain and its outer domain, the surrounding grid will overlap with the background grid of the overall flow domain, so the grid cutting is performed on the boundary of the outer domain of the rocket 3 to cut off the repeated grid in the background domain, which is also called "hole digging". The contact between the two sides of the cutting boundary is established by interpolation fitting to establish the contact of the flow parameters. When the rocket 3 moves, the cutting and hole digging of the new position grid are performed.

[0071] S4.2, establish the rocket motion equation:

[0072] According to the rocket motion equation, the motion velocity of the rocket is calculated, so that the rocket flies at the actual velocity and trajectory, thereby simulating the rocket launching process.

[0073] S4.3, define the flow parameters of the rocket engine combustion chamber (wherein the pressure increases with time), and define and calculate the atmospheric environment parameters according to the atmospheric environment during launching.

[0074] Go to step 5.

[0075] Step 5, solve the flow model of the above rocket take-off thermal environment and its boundary parameters, the flow parameters of the rocket engine combustion chamber, and the atmospheric environment parameters to obtain the flow field results: use Realizable k-ε turbulence model to solve the viscous dissipation in N-S equation, and use central difference to disperse the viscous flux, complete the N-S equation solution, and obtain the Mach number, temperature, pressure flow field atlas, and the load change curve of the gas jet impacting the launching cylinder.

[0076] S5.1, establish Realizable k-ε turbulence model

[0077] The turbulence model is established to close the Reynolds stress. Currently, the commonly used turbulence models include zero equation model, one equation model, standard k-ε two equation model, and Realizable k-ε model. Since the two equation model has higher calculation precision, it is more and more widely used in engineering calculation. Turbulence kinetic energy k and dissipation rate ε are usually used to describe the characteristic scale of turbulence. Realizable k-ε model is improved on the basis of standard k-ε equation, and is more suitable for the rocket gas jet impact problem studied by the method. It can better keep the Reynolds stress consistent with the real turbulence, and accurately simulate the diffusion rate of planar jet. The generation term in the turbulence dissipation rate ε equation no longer includes the generation term G k in the turbulence kinetic energy k equation, and the coefficient C t in the turbulence viscosity μ μ is not a constant, but is related to the strain rate. Such form better represents energy conversion.

[0078] The turbulence kinetic energy k equation is expressed as:

[0079]

[0080] In the formula, μ is the viscosity, μ i is the viscosity of component i, G k is the generation term of turbulence kinetic energy k caused by the average velocity gradient, the constant coefficient σ k = 1.0, x i , x j are the displacement amounts of the gas in X and Y directions, and the turbulence dissipation rate ε equation is:

[0081]

[0082] wherein the first constant coefficient σ ε = 1.2, the second constant coefficient C1 = 1.44, the third constant coefficient C2 = 1.9, I is the average characteristic strain rate, v is the velocity of the rocket gas flow, p is the density of the rocket gas, p is the pressure, and t is the time.

[0083] S5.2, the simulation results are obtained after solving the viscous dissipation in the N-S equation of the established RNG k-epsilon turbulent flow model, and the simulation results include the Mach number, temperature, pressure flow field atlas, and the load change curve of the gas jet impacting the launch cylinder.

[0084] Embodiment

[0085] A method for simulating the flow field of a rocket gas jet impacting a vertical launching device, according to the steps in the above specific embodiment, comprising the following steps:

[0086] Step 1, establishing a three-dimensional model of the vehicle-mounted vertical launching device;

[0087] In combination with Figure 2 The three-dimensional model includes a rocket 2, a vehicle body including tires, a vehicle head, a chassis, a cabin 3, and a launch cylinder; and a 1:1 isometric modeling is performed on the actual engineering launch vehicle device.

[0088] Step 2, first, the three-dimensional model of the vehicle-mounted vertical launching device is divided into flow regions, and the flow regions (i.e., the overall calculation domain) are processed by block grid generation strategy, i.e., the overall calculation domain is divided into four sub-domains, which are the rocket motion flow domain, the launch cylinder and the upper flow domain, the vehicle body and the jet aggregation flow domain around the vehicle body, and the remaining flow domain, and the calculation structure grid is generated for each sub-domain. The grid area of the rocket motion flow domain is the main grid, the grid area of the remaining flow domain and the grid area of the launch cylinder and the upper flow domain are background grids, and the main grid is connected with the two background grids to form a nested grid.

[0089] In combination with Figure 3 The total number of grids of the generated vehicle-mounted vertical launching device and the rocket motion domain is about 17.1 million hexahedral units.

[0090] Step 3, the gas components include sp1, sp2, and sp3, and the molar mass percentages are 35%, 24%, and 41% respectively, which are substituted into the multi-component equation for solving. The multi-component N-S equation containing gas transport is discretized into an improved Roe finite volume format, the positive factor value is determined to be 0.2 according to the properties of the gas components, and the viscous flux at the cell boundary of the calculation structure grid is obtained by numerically integrating the differential equation of the flow equation.

[0091] Step 4, establishing a nested grid model, specifically including cutting, hole digging and interpolation, forming a motion boundary, solving the rocket motion equation, and defining the engine ignition pressure distribution;

[0092] The calculation is from the rocket engine ignition to the motion distance of 17.5km; the environmental temperature is 300k; the initial speed when the rocket main engine is ignited is 26.046m / s; the external atmospheric environment pressure is 101325Pa

[0093] The engine internal working pressure changes with time as shown in Figure 9 The engine internal temperature during operation is 3103k

[0094] Step 5, use Realizable k-ε turbulence model to solve the viscosity flux in the above model N-S equation, output the Mach number, temperature, pressure flow field and the load change curve of the gas impact launch cylinder.

[0095] Using Realizable k-ε turbulence model for solving and calculating, the rocket gas jet impact vertical launch device flow field simulation method designed in this paper takes about 15 days to calculate in the case of using 64 CPU core numbers for parallel calculation, and the output Mach number field, temperature field, pressure field atlas, and launch device stress curve are obtained, as shown in Figures 5-8 , Figure 5 The maximum Mach number of the rocket flow field is about 4.4, Figure 6 The maximum temperature of the rocket during launch is located at the first Mach disc (normal shock) wave front position downstream of the nozzle, Figure 7 The external pressure peaks of the rocket are located at the first, second, third, … Mach disc of the nozzle, and show a gradually decaying trend, Figure 8 The vertical (along the z-axis) stress change of the launch cylinder is shown, and the stress reaches a peak value when the rocket engine works for about 0.1s. The method of the present application improves the calculation accuracy while saving the calculation time, and better simulates the influence of the gas jet impact effect of the rocket ignition during the vehicle-mounted vertical launch on the launch device, and the calculation result provides an important analysis means for the lightweight design and safety of the vehicle-mounted vertical launch device.

Claims

1. A method for simulating the flow field of a vertical launch device impacted by a rocket gas jet, characterized in that, The steps are as follows: Step 1, a three-dimensional model of the vehicle-mounted vertical launching device is established, and the three-dimensional model includes a vehicle body (1), a launching cylinder (2), and a rocket (3), and proceeds to Step 2; Step 2, the established three-dimensional model of the vehicle-mounted vertical launching device is divided into flow regions, and a moving nested grid system is constructed: The established three-dimensional model of the vehicle-mounted vertical launching device is divided into flow regions, and a moving nested grid system is constructed by using a block grid generation strategy, i.e., the entire flow region is first divided into multiple sub-regions, and then a calculation structure grid is generated for each sub-region, and proceeds to Step 3; Step 3, a theoretical model is selected according to the moving nested grid system, and a flow model of the rocket take-off force and heat environment is established: The multi-component N-S equation containing gas transport is discretized into an improved Roe finite volume format, the positive factor value is determined, and the viscous flux at the unit boundary of the calculation structure grid is obtained by numerically integrating the differential unit flow equation according to the improved Roe finite volume format; The flow model of the rocket take-off force and heat environment is selected according to the moving nested grid system and is established as shown in equations (1) to (9), and the details are as follows: S3.1, before establishing the flow model of the rocket take-off force and heat environment, the following basic assumptions are made for the gas jet: 1) the gas jet satisfies the continuous medium; 2) the gas jet is a compressible pure gas phase medium; 3) no chemical reaction occurs inside the gas jet; 4) a gas and air multi-component mixed flow model is used, and the multiple components all satisfy the ideal gas state equation; S3.2, the above assumptions ensure that the gas is a Newtonian viscous fluid, so the gas multi-component transport equation is established as where Y a is the mass fraction of the rocket fuel component a, is the flux of a physical quantity, such as is the diffusion term for the rocket fuel component a, and is the flux of the diffusion of the rocket fuel component a, R a is the net production rate of the rocket fuel component a by chemical reaction, S a is the production rate caused by the rocket fuel dispersion phase from the self-defined source term, t is the flight time of the multi-nozzle rocket, p is the density of the rocket fuel, v is the flow velocity of the rocket fuel, and N is the number of rocket fuel components. In equation (1), the flux of component diffusion is In formula (2), D a,m D is the diffusion coefficient of component a in the mixture T,a is the divergence of the gas flow element, and T is the temperature of the rocket gas S3.3, the above assumption that the gas is a compressible gas is made, and in the rectangular coordinate system, the conservation form of the N-S equation is established as wherein V is a velocity vector of the rocket flight, S is a gas flow cross-sectional area, Q is a rocket gas combustion heat, is a generalized source term of the multi-nozzle rocket, are direction vectors in X, Y, Z directions of the rectangular coordinate system, E, F, and G are rocket gas flow flux vectors in X, Y, Z directions of the rectangular coordinate system, E v , F v , and G v are rocket gas viscous flux vectors in X, Y, Z directions of the rectangular coordinate system; S3.4, the N-S equation is discretized by the improved Roe format to obtain the viscous flux at the unit boundary of the calculation structure grid: The viscous flux at the unit boundary of the calculation structure grid is expanded as shown below by using the basic form of the Roe format: ξ = |U| (5) where, is the Roe-form numerical inviscid flux, β is the positive factor, which is between 0-0.5 according to different gas physical properties; Δ(*) is the change of physical quantity *; ξ is the basic upwind dissipation term, δp u is the interface correction pressure driven by velocity gradient, δp p is the interface correction pressure driven by pressure gradient, δU u is the interface correction velocity driven by velocity gradient, δU p is the interface correction velocity driven by pressure gradient, n x , n y , n z are the components of the interface normal vector on X, Y, Z axes respectively; c is the sound speed of the atmosphere, U is the interface normal velocity; ρ is the density of the rocket gas, u, b, w are the rocket gas flow velocities in X, Y, Z directions respectively, p is the pressure, B is the total energy, H is the total enthalpy of the rocket gas; proceeds to Step 4; Step 4, boundary parameters of the flow model of the rocket take-off force and heat environment are set: The background grid and the moving grid of the moving nested grid system are defined, and the background grid and the moving grid are respectively cut and drilled to construct the flow model of the moving boundary, the flow model of the rocket take-off force and heat environment and the rocket motion equation are solved, the flow parameters of the rocket engine combustion chamber are defined, and the atmospheric environment parameters are calculated according to the atmospheric environment during launching, and proceeds to Step 5; Step 5, the flow field results are obtained by solving the above flow model of the rocket take-off force and heat environment, its boundary parameters, the flow parameters of the rocket engine combustion chamber, and the atmospheric environment parameters: the Realizable k-ε turbulence model is used to solve the viscous dissipation in the N-S equation, the central difference is used to discretize the viscous flux, the entire N-S equation solution is completed, and the Mach number, temperature, and pressure flow field graphs, as well as the load variation curve of the gas jet impacting the launching cylinder, are obtained.

2. The method of claim 1, wherein: In Step 1, the three-dimensional model of the vehicle-mounted vertical launching device is established, and the details are as follows: The three-dimensional model comprises a vehicle body (1), a launching cylinder (2) and a rocket (3), the vehicle body (1) comprises a cabin body, a vehicle head, a tire and a chassis; the model parameters of the vehicle body (1), the launching cylinder (2) and the rocket (3) are consistent with the actual engineering, and the vehicle body (1) and the rocket (3) are modeled in a 1:1 equal ratio with the actual engineering launching vehicle and rocket.

3. The method of claim 2, wherein: The three-dimensional model parameters comprise the following: the geometric size of the vehicle body (1), the diameter, length and throat radius parameters of the rocket (3), the inner diameter, outer diameter and length of the launching cylinder (2).

4. The method of claim 1, wherein: In step 2, the established three-dimensional model of the vehicle-mounted vertical launching device is divided into a flow region, and a moving nested grid system is constructed, as follows: The established three-dimensional model of the vehicle-mounted vertical launching device is divided into a flow region, and a moving nested grid system is constructed, as follows:

5. The method of claim 4, wherein: Firstly, the established three-dimensional model of the vehicle-mounted vertical launching device is divided into a flow region, and a moving nested grid system is constructed, as follows:

6. The method of claim 5, wherein: The flow region is divided into four sub-regions, i.e., a rocket movement flow region, a launching cylinder and upper flow region, a vehicle body and vehicle body surrounding jet flow aggregation flow region, and a remaining flow region, and structured grid is generated for each sub-region; the grid region of the rocket movement flow region is the main grid, the grid region of the remaining flow region and the grid region of the launching cylinder and upper flow region are background grids, and the main grid is connected with the two background grids to solve the movement boundary problem of the rocket taking off, thereby constructing the moving nested grid system.

7. The method of claim 1, wherein, The launching cylinder flow region and the vehicle body and vehicle body surrounding jet flow aggregation flow region are subjected to structured grid encryption processing to ensure the orthogonality and smoothness of the grid, and the grid in the fluid region far away from the vehicle-mounted vertical launching device is transitioned from dense to sparse. In step 4, the boundary parameters of the flow model of the rocket taking-off thermal environment are set, as follows: S4.1, setting nested cutting boundaries and background grids: When the rocket movement flow region and its outer region move, the surrounding grid will overlap with the background grid of the whole flow region, so the grid cutting is performed with the outer region around the rocket (3) as the boundary, the repeated grid in the background region is cut off, which is also called "hollowing", and the contact between the two sides of the cutting boundary is established by interpolation fitting; when the rocket (3) moves, the cutting and hollowing of the new position grid are performed; S4.2, establishing a rocket movement equation: According to the rocket movement equation, the movement speed of the rocket is calculated, so that the rocket flies at the actual speed and trajectory, thereby simulating the rocket launching process; 8. The method of claim 1, wherein, S4.3, defining the flow parameters of the rocket engine combustion chamber, and defining and calculating the atmospheric environment parameters according to the atmospheric environment during launching. The pressure in the flow parameters of the rocket engine combustion chamber increases with time.

9. The method of claim 1, wherein, In step 5, the flow model of the above rocket launch force-thermal environment and its boundary parameters, the rocket engine combustion chamber flow parameters and the atmospheric environment parameters are solved to obtain the flow field results: the Realizable k-ε turbulence model is used to solve the viscous dissipation in the N-S equation, the central difference is used to disperse the viscous flux, the entire N-S equation solution is completed, the Mach number, temperature, pressure flow field atlas, and the load change curve of the gas jet impacting the launch cylinder are obtained, which are as follows: S5.1, establish Realizable k-ε turbulence model The turbulence model kinetic energy k equation is expressed as: where μ is the viscosity, μ i is the viscosity of component i, k represents the turbulent kinetic energy, μ t represents the turbulent viscosity, C μ represents the coefficient in the turbulent viscosity, G k is the production term of the turbulent kinetic energy k caused by the average velocity gradient, the constant coefficient σ k = 1.0, x i , x j is the displacement amount of the fuel gas in the X, Y directions, and the turbulent dissipation rate ε equation is: wherein the first constant coefficient σ ε = 1.2, the second constant coefficient C1 = 1.44, the third constant coefficient C2 = 1.9, I is the average characteristic strain rate, v is the velocity of the rocket gas flow, p is the density of the rocket gas, p is the pressure, and t is the time. S5.2, after solving the viscous dissipation in the N-S equation for the established RNG k-ε turbulence model, the simulation results are obtained, including the Mach number, temperature, pressure flow field atlas, and the load change curve of the gas jet impacting the launch cylinder.

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