Aerodynamic heating and thermal response coupling calculation method considering ablation pyrolysis products injection

Abstract

The application discloses a method for calculating aerodynamic heat and thermal response coupling considering ablation pyrolysis product injection, relates to the field of aerodynamic heat and thermal protection coupling numerical calculation, and establishes the following methods: a numerical calculation method for aerodynamic heat environment under the condition of mass injection; a numerical calculation method for thermal response considering surface ablation reaction, internal pyrolysis reaction, pyrolysis gas flow and solid heat conduction; and a method for coupling data transmission between the two methods. The method can simulate the generation, transportation and injection process of ablation pyrolysis thermal protection system surface ablation reaction products and internal pyrolysis reaction products of high-speed aircraft, and can calculate the coupling aerodynamic heat / thermal response result of the ablation pyrolysis thermal protection system of high-speed aircraft considering mass injection.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of numerical calculation of aerodynamic heating and thermal protection coupling, in particular, to an aerodynamic heating and thermal response coupling calculation method considering ablation pyrolysis product injection. BACKGROUND

[0002] The present application belongs to the field of aerodynamic heating and thermal protection technology, and flow-thermal-solid coupling calculation, and can be coupled to simulate the aerodynamic heating environment and material thermal response of a high-speed (Mach number greater than 5) vehicle thermal protection system surface under the condition of ablation and internal pyrolysis gas injection. When a vehicle flies at high speed in the atmosphere, it will face a severe aerodynamic heating environment. In order to ensure the thermal safety of the vehicle, an ablation thermal protection system is often used. Ablation thermal protection systems have been used for a long time and have been widely used in various high-speed (Mach number greater than 5) vehicles such as ballistic missiles, return capsules / entry vehicles, near-space vehicles, etc. The working principle of the ablation thermal protection system is that the material will undergo ablation and pyrolysis under severe aerodynamic heating environment, thereby causing mass loss of the material, absorbing and carrying away a large amount of heat, and preventing part of the heat from being transferred into the vehicle. The internal pyrolysis, surface ablation reaction, melting, evaporation / sublimation and other processes in the ablation thermal protection system all involve the surface mass injection phenomenon of the fluid leaving the material surface and entering the flow field. There is an interaction between the aerodynamic heating environment and the structural thermal response of a high-speed (Mach number greater than 5) vehicle, especially when the fluid domain of the chemical non-equilibrium incoming flow, the solid domain of the ablation pyrolysis material and the mass injection behavior at the gas-solid interface are considered, which will lead to a more close and complex aerodynamic heating / thermal response coupling effect.

[0003] The existing simulation methods for aerodynamic heating / thermal response considering ablation pyrolysis product injection have certain limitations. In terms of flow field simulation, some methods ignore the influence of mass injection on the thermal environment and only set temperature boundary conditions. Some methods modify the aerodynamic heating environment without considering mass injection by using a mass injection coefficient in order to simplify the calculation. The above methods do not calculate the behavior of the fluid entering the high-speed flow field from the surface, although the calculation cost is low, but it is difficult to simulate the overall influence of the mass injection phenomenon on the aerodynamic heating environment. In addition, when the vehicle is in a high-enthalpy flow, chemical reactions will occur in the boundary layer. After the mass injection gas enters the boundary layer, it may react with these gases again, thereby affecting the aerodynamic heating environment. Currently, there are few methods to simulate this phenomenon.

[0004] In terms of solid domain simulation, it mainly involves the simulation of solid heat transfer, porous medium flow and heat transfer, ablation pyrolysis thermal response, etc. These processes will affect the generation and transport of mass injection gas and determine the conditions of mass injection gas, but the existing methods often ignore or simplify the simulation of the above processes.

[0005] The existing method causes a lack of effective evaluation means on the forefront engineering problems of accurate prediction of the thermal environment of the aircraft and evaluation and design of the heat protection capability, so that the thermal safety of a new high-speed (Mach number greater than 5) aircraft is difficult to guarantee. SUMMARY

[0006] The application aims to provide an aerodynamic heat / thermal response coupling calculation method considering the injection of ablation pyrolysis products, which is currently lacking, and provide support for the design evaluation of the thermal protection system of a high-speed (Mach number greater than 5) aircraft.

[0007] To achieve the above application purpose, the application provides an aerodynamic heat / thermal response coupling calculation method considering the injection of ablation pyrolysis products, which comprises the following steps:

[0008] Step 1: determining the geometric shape of the aircraft, and drawing a flow field grid and a structure field grid;

[0009] Step 2: selecting a flight trajectory point, and based on the flight Mach number, height, attack angle and sideslip angle of the selected flight trajectory point at the corresponding time, performing aerodynamic heat environment solving based on the flow field grid to obtain aerodynamic heat environment calculation results;

[0010] Step 3: interacting the flow field data and the structure field data at the wall surface position, and transferring the wall surface heat flow distribution and pressure distribution in the aerodynamic heat environment calculation results to the structure field surface grid points to obtain boundary conditions for thermal response calculation;

[0011] Step 4: establishing a mass conservation equation considering the ablation reaction of the surface of the aircraft, the internal pyrolysis and the flow of the pyrolysis gas, establishing an energy conservation equation considering the endothermic reaction of the ablation reaction of the surface of the aircraft and the internal pyrolysis reaction, the heat conduction of the solid phase and the convective heat transfer of the pyrolysis gas, and based on the boundary conditions for thermal response calculation obtained in step 3, performing thermal response calculation by using the mass conservation equation and the energy conservation equation, simulating for a set time length, and obtaining the temperature distribution of the aircraft structure, the ablation reaction of the surface of the aircraft and the internal pyrolysis reaction and the flow of the pyrolysis gas of the aircraft;

[0012] Step 5: interacting the flow field data and the structure field data at the wall surface position, obtaining the surface temperature distribution of the aircraft based on the temperature distribution of the aircraft structure, obtaining the ablation reaction product components and the injection mass flow rate distribution of the surface of the aircraft based on the ablation reaction of the surface of the aircraft, obtaining the internal pyrolysis reaction product components and the injection mass flow rate distribution at the surface of the aircraft based on the internal pyrolysis reaction and the flow of the pyrolysis gas of the aircraft, and transferring the surface temperature distribution of the aircraft, the ablation reaction product components and the injection mass flow rate distribution of the surface of the aircraft and the internal pyrolysis reaction product components and the injection mass flow rate distribution at the surface of the aircraft to the flow field wall surface grid points to obtain boundary conditions for aerodynamic heat environment calculation;

[0013] Step six: determining the time according to the set time length simulated in step four, determining the flight trajectory based on the time, determining the flight Mach number, height, angle of attack and side slip angle based on the flight trajectory, inputting the aircraft surface temperature distribution, aircraft surface ablation reaction product component and injection mass flow rate distribution and the aircraft surface internal pyrolysis reaction product component and injection mass flow rate distribution transferred in step five into the mass injection momentum boundary condition, performing aerodynamic thermal environment calculation based on the flow field grid, and obtaining the calculation aerodynamic thermal environment result;

[0014] Step seven: determining whether the flight trajectory is calculated, if yes, ending; if not, returning to step two and continuing to perform aerodynamic thermal environment solving along the flight trajectory point.

[0015] Wherein, since the aerodynamic thermal environment of the aircraft and the structural thermal response interact with each other, and the mass injection behavior will have an important influence on the aerodynamic thermal environment, resulting in the change of the aerodynamic heat / force load received by the material, the temperature distribution of the material and the flow of the porous medium will change, and the generation, conditions and transport process of the ablation pyrolysis gas depend on the change of the above conditions, so the influence of mass injection on the aerodynamic thermal / thermal response coupling thermal effect of the ablation heat protection system cannot be ignored. For coupling simulation, on the one hand, the key processes of each physical field and the key parameter transmission between each physical field should be fully considered; on the other hand, the time scale of each physical field simulation and the data exchange frequency between each physical field should also be considered.

[0016] Therefore, the application designs an ablation pyrolysis product injection aerodynamic thermal / thermal response coupling calculation method considering the aerodynamic thermal environment calculation under the condition of mass injection; the thermal response calculation considering the surface ablation reaction, internal pyrolysis reaction and pyrolysis gas flow, solid heat conduction; and the coupling data transmission method strategy between the two.

[0017] The coupling calculation method provided by the application can describe the process that the gas generated by the pyrolysis material under heat reaction flows through the porous medium to the surface and the process that the gas generated by the ablation material surface under heat reaction; and the process that these gases are injected into the flow field with a certain mass flow rate and interact with the main flow to change the behavior of the flow field; the coupling calculation method can obtain the coupling aerodynamic thermal / thermal response result of the ablation pyrolysis heat protection system of the high-speed (Mach number greater than 5) aircraft considering mass injection.

[0018] Preferably, the step three utilizes the inverse distance weighting method to interpolate between the spatial flow field grid and the solid field grid, and transfers the wall heat flux distribution and pressure distribution in the aerodynamic thermal environment calculation result to the structure field surface grid point to obtain the boundary condition of the thermal response calculation; the step five is specifically: the flow field data and the structure field data are interacted at the wall position, and the aircraft structure temperature distribution, the aircraft surface ablation reaction, and the aircraft internal pyrolysis reaction and pyrolysis gas flow obtained by the step four are transferred to the flow field wall grid point by the inverse distance weighting method to obtain the boundary condition of the aerodynamic thermal environment calculation;

[0019] The inverse distance weighting method calculation formula is:

[0020] ;

[0021] ;

[0022] ;

[0023] wherein, is the physical quantity on the known interpolation point, is the physical quantity on the to-be-interpolated point, is the distance between the known interpolation point and the to-be-interpolated point, and are attenuation control quantities, is the weight value of the interpolation node i, which represents the contribution of the physical information at the point to the to-be-interpolated point at the time interval, n is the number of interpolation points, x and y are the horizontal coordinate and vertical coordinate of the to-be-interpolated point, respectively, x i and y i are the horizontal coordinate and vertical coordinate of the known interpolation point, respectively.

[0024] Preferably, the step two is specifically: selecting a flight trajectory point, based on the flight Mach number, height, attack angle and side slip angle of the selected flight trajectory point at the corresponding time, the aerodynamic thermal environment calculation result is obtained by solving the three-dimensional compressible Navier-Stokes equation based on the flow field grid by using the finite volume method; the step six utilizes the finite volume method to solve the three-dimensional compressible Navier-Stokes equation, and performs aerodynamic thermal environment calculation based on the flow field grid to obtain the calculation aerodynamic thermal environment result.

[0025] Preferably, the mass conservation equation is:

[0026] ;

[0027] wherein, is the material porosity, is the pyrolysis gas constant, is the temperature, is the pressure, t is time, p is the material permeability, v is the pyrolysis gas velocity, p0 is the material porosity before pyrolysis, p0 is the material density before pyrolysis, p is the material porosity after pyrolysis, p is the material density after pyrolysis, p is the material solid density, D is the partial derivative operator.

[0028] The energy conservation equation is preferably:

[0029] ;

[0030] where, p is the pyrolysis gas density, cp is the gas specific heat at constant pressure, T is the temperature, t is time, p is the material porosity, p is the material solid density, cp is the solid specific heat at constant pressure, k is the thermal conductivity, D is the Darcy velocity of the gas, h is the gas product enthalpy, h is the material enthalpy, q is the pyrolysis heat flux.

[0031] The mass ejection momentum boundary condition is preferably:

[0032] ;

[0033] ;

[0034] ;

[0035] ;

[0036] ;

[0037] ;

[0038] where, T is the temperature at the surface, i.e. the temperature of the surface mass ejection gas, T is the set temperature (in the calculation, the surface temperature distribution calculated in step four and transferred in step five), and v is the velocity of the surface in two directions not normal to the surface, density at the surface, pressure at the first layer of grid near the surface, density at the first layer of grid near the surface, velocity at the first layer of grid near the surface, mass injection gas constant, mass flow rate of the surface mass injection gas, velocity at the surface.

[0039] The one or more technical solutions provided by the present application have at least the following technical effects or advantages:

[0040] The present application establishes an aerodynamic heating environment numerical calculation method under the condition of mass injection; a thermal response numerical calculation method considering surface ablation reaction, internal pyrolysis reaction and pyrolysis gas flow, solid heat conduction; and an aerodynamic heating / thermal response coupling calculation method of ablation pyrolysis product injection considering the coupling data transmission method strategy. The generation, transportation and injection process of the surface ablation reaction product and the internal pyrolysis reaction product of the ablation pyrolysis thermal protection system of the high-speed (Mach number greater than 5) aircraft can be simulated, the coupling aerodynamic heating / thermal response result of the ablation pyrolysis thermal protection system of the high-speed (Mach number greater than 5) aircraft considering mass injection can be calculated, and the evaluation and design of the ablation pyrolysis thermal protection system of the high-speed (Mach number greater than 5) aircraft can be supported. BRIEF DESCRIPTION OF DRAWINGS

[0041] The accompanying drawings, which are included to provide a further understanding of the embodiments of the present application and constitute a part of the present application, do not limit the embodiments of the present application;

[0042] Figure 1 FIG. 1 is a flowchart of an aerodynamic heating and thermal response coupling calculation method considering ablation pyrolysis product injection;

[0043] Figure 2 FIG. 2 is a coupling data transmission schematic diagram;

[0044] Figure 3 FIG. 3 is a schematic diagram of the geometry and size of the calculation object;

[0045] Figure 4 FIG. 4 is a flow field calculation grid schematic diagram;

[0046] Figure 5 FIG. 5 is a structure field calculation grid schematic diagram;

[0047] Figure 6 FIG. 6 is a structure temperature distribution schematic diagram at time 1s;

[0048] Figure 7 FIG. 7 is a structure temperature distribution schematic diagram at time 2s;

[0049] Figure 8 is the schematic diagram of the flow field temperature distribution and the flow line near the wall surface at the time of 1s;

[0050] Figure 9 is the schematic diagram of the wall surface heat flow at the time of 0s, 1s and 2s considering / without considering the mass injection. DETAILED DESCRIPTION

[0051] In order to enable the above-mentioned objects, features and advantages of the present application to be more clearly understood, the following will further describe the present application with reference to the accompanying drawings and specific embodiments. It should be noted that the embodiments of the present application and the features in the embodiments can be combined with each other without conflict.

[0052] In the following description, many specific details are set forth in order to provide a thorough understanding of the present application, however, the present application can be practiced in other manners different from those described herein within the scope of the present application, therefore, the scope of protection of the present application is not limited by the specific embodiments disclosed below.

[0053] Embodiment one;

[0054] Please refer to Figures 1-2 , Figure 1 is the schematic diagram of the flow field temperature distribution and the flow line near the wall surface at the time of 1s; Figure 2 is the schematic diagram of the flow field temperature distribution and the flow line near the wall surface at the time of 1s; Figure 2 wherein p is pressure, q is heat flow, T is temperature, c is component, subscript w is wall surface, subscript py is pyrolysis gas, and subscript ab is ablation gas. The present application provides a method for calculating the aerodynamic heat / thermal response coupling considering the ablation pyrolysis product injection, which calculates the flow field thermal environment by solving the three-dimensional compressible Navier-Stokes equation, calculates the structure temperature field by solving the three-dimensional heat conduction equation, simulates the mass injection gas generation, transportation and reaction process by adding the multi-component gas mass injection momentum boundary condition, high-temperature gas chemical reaction model, porous medium Darcy law and ablation pyrolysis reaction model, and performs coupling solution through the established field data transmission method.

[0055] The method for calculating the aerodynamic heat / thermal response coupling considering the ablation pyrolysis product injection has the following specific steps:

[0056] Step one: determine the geometric shape of the calculation object, draw the flow field grid and the structure field grid, and the calculation object is usually an aircraft.

[0057] Step two: select the flight trajectory point, according to the flight Mach number, height, angle of attack, side slip angle at this time (the corresponding time of the selected flight trajectory point), the aerodynamic heat environment is solved based on the flow field grid; when solving this step, the wall surface is set as a viscous solid wall boundary condition, and the wall surface temperature is selected according to the need. The three-dimensional compressible Navier-Stokes equation is solved by using the finite volume method

[0058] The aerodynamic heat environment calculation result under this condition is obtained. The three-dimensional compressible Navier-Stokes equation can refer to Gui Yewei, Tang Wei, Du Yanxia. Near space hypersonic vehicle thermal safety [M]. Beijing: National Defense Industry Press, 2019: p24-26. The embodiments of the application do not make corresponding redundant description.

[0059] Step three: the flow field data and the structure field data are interacted at the wall surface position, the wall heat flow distribution and the pressure distribution in the aerodynamic heat environment calculation result are transmitted to the structure field surface grid point, the boundary conditions for heat response calculation, the inverse distance weighting method is used to interpolate between the flow field grid and the solid field grid in space, and the inverse distance weighting method calculation formula is as follows:

[0060] ;

[0061] ;

[0062] ;

[0063] Among them, is the physical quantity on the known interpolation point, is the physical quantity on the to-be-interpolated point, is the distance between the known interpolation point and the to-be-interpolated point, and are attenuation control quantities, generally , . is the weight of the interpolation node i, which represents the contribution of the physical information at this point to the to-be-interpolated point in the time interval. n is the number of interpolation points, x and y are the horizontal and vertical coordinates of the to-be-interpolated point, respectively, x i and y i are the horizontal and vertical coordinates of the known interpolation point, respectively.

[0064] Step four: considering the surface ablation reaction, internal pyrolysis and pyrolysis gas flow, the mass conservation equation can be established:

[0065] ;

[0066] Considering the heat absorption of the surface ablation reaction and the internal pyrolysis reaction, the heat conduction of the solid phase, and the convective heat transfer of the pyrolysis gas, the energy conservation equation can be established:

[0067] ;

[0068] The heat flow distribution and pressure distribution passed in step three are taken as input boundary conditions, and heat response calculation is carried out based on the structure field grid and material properties by using the mass conservation equation and the energy conservation equation, and the temperature rise inside the structure, the surface ablation reaction, the internal pyrolysis reaction and the pyrolysis gas flow are simulated for a certain time length (which can be selected according to the situation).

[0069] Step five: The flow field data and the structure field data are interacted at the wall position, and the surface temperature distribution, the surface ablation reaction product component and the injection mass flow rate distribution, the internal pyrolysis reaction product component and the injection mass flow rate distribution at the surface in the heat response calculation result in step four are passed to the flow field wall grid points by the inverse distance weighting method in step three, and are taken as boundary conditions for aerodynamic thermal environment calculation.

[0070] Step six: According to the time length simulated in step four, the flight trajectory is determined, the flight Mach number and the height are determined, and the surface temperature distribution, the surface ablation reaction product component and the injection mass flow rate distribution, the internal pyrolysis reaction product component and the injection mass flow rate at the surface passed in step four are taken as mass injection momentum boundary conditions, and the three-dimensional compressible Navier-Stokes equation in step one is solved by using the finite volume method, and aerodynamic thermal environment calculation is carried out based on the flow field grid, and the calculation result is obtained. The mass injection momentum boundary condition is:

[0071] ;

[0072] ;

[0073] ;

[0074] ;

[0075] ;

[0076] ;

[0077] The injection conditions of different component gases are respectively substituted into the above formula.

[0078] Step seven: It is judged whether the flight trajectory is calculated or not, if yes, the process is ended, if not, the process returns to step two, and the aerodynamic thermal environment solution is continued along the flight trajectory point.

[0079] In step four, the thermal response calculation is performed using the mass conservation equation and the energy conservation equation. The simulation is set for a certain duration to obtain the temperature distribution of the aircraft structure, the ablation reaction on the aircraft surface, and the pyrolysis reaction and pyrolysis gas flow inside the aircraft. For specific implementation methods, please refer to Liu Xiao. Research on finite element calculation method of three-dimensional thermal response of ablation thermal protection system [D]. Mianyang: China Aerodynamics Research and Development Center, 2017. and Mark E. Multidimensional Modeling of Ablation Heat Transfer [J]. Journal of Thermophysics and HeatTransfer, 2023, 37(1):104-118. The embodiments of this invention will not be described in detail. This invention establishes a numerical calculation method for aerodynamic thermal environment under mass ejection conditions; a numerical calculation method for thermal response considering surface ablation reaction, internal pyrolysis reaction and pyrolysis gas flow, and solid heat conduction; and a coupled data transfer method strategy for aerodynamic thermal / thermal response coupling calculation method considering the ejection of ablation pyrolysis products. It can simulate the generation, transport, and ejection processes of surface ablation reaction products and internal pyrolysis reaction products of ablation pyrolysis thermal protection systems for high-speed (Mach number greater than 5) aircraft, and can calculate the coupled aerodynamic / thermal response results of ablation pyrolysis thermal protection systems for high-speed (Mach number greater than 5) aircraft considering mass ejection.

[0080] Example 2;

[0081] Based on Example 1, Example 2 illustrates the method with specific examples:

[0082] Step 1: Determine the geometric shape of the object to be calculated, such as... Figure 3 As shown; the flow field mesh and structure field mesh are plotted as follows: Figure 4 and Figure 5 As shown.

[0083] Step Two: Select the flight trajectory point. In this embodiment, time 0s is selected, with a flight Mach number of 10, a flight altitude of 55km, an angle of attack of 0°, and a sideslip angle of 0°. Figure 4 The aerodynamic thermal environment is solved using the flow field mesh; in this step, the wall is set as a viscous solid wall boundary condition, and the wall temperature is set to 300°C. The aerothermal environment calculation results under this condition were obtained by solving the three-dimensional compressible Navier-Stokes equations using the finite volume method, and the file FlowCalculate_0s.dat was generated.

[0084] Step 3: Calculate the values ​​in the file FlowCalculate_0s.dat generated in Step 2 based on the flow field mesh. Figure 4) and structural field mesh ( Figure 5 The inverse distance weighting method is used to calculate the aerodynamic thermal environment at the gas-solid interface. The calculation results are interpolated from the flow field grid points to the structural field grid points, generating the file Flow-Solid_0s.dat. This file contains the wall heat flux distribution and pressure distribution calculated in step two. The inverse distance weighting method calculation formula is as follows:

[0085] ;

[0086] ;

[0087] ;

[0088] Step 4: Using the wall heat flux distribution and pressure distribution calculated in Flow-Solid_0s.dat in Step 3 as input boundary conditions, based on... Figure 5 Based on the structural field mesh and material properties, a 1-second thermal response calculation was performed to obtain the internal temperature rise, surface ablation reaction, internal pyrolysis reaction, and pyrolysis gas flow. The material property parameters used are shown in Tables 1 and 2. Table 1 shows the material properties before pyrolysis, and Table 2 shows the material properties after complete pyrolysis. The thermal response calculation employed the mass conservation equation and the energy conservation equation.

[0089] ;

[0090] ;

[0091] The thermal response calculation results under these conditions are obtained, and the file SolidCalculate_1s.dat is generated.

[0092] Table 1 Material properties before pyrolysis

[0093]

[0094] Table 2 Material properties after pyrolysis

[0095]

[0096] Step 5: Calculate the values ​​in the file FlowCalculate_1s.dat generated in Step 4 based on the flow field mesh. Figure 4 ) and structural field mesh ( Figure 5) and the thermal response results at the gas-solid interface are interpolated from the structural field grid points to the flow field grid points by inverse distance weight method to generate the file Solid-Flow _1s.dat, which contains the surface temperature distribution, the surface ablation reaction product composition and the injection mass flow rate distribution, the surface internal pyrolysis reaction product composition and the injection mass flow rate distribution calculated in step four, and the inverse distance weight method calculation formula is shown in step three.

[0097] Step six: according to the simulation time of 1s in step three, the flight time is determined to be 1s, and the flight trajectory is determined to be 1s, the flight Mach number is 11, the flight height is 50km, the attack angle is 5°, and the sideslip angle is 0°; the surface temperature distribution, the surface ablation reaction product composition and the injection mass flow rate distribution, the surface internal pyrolysis reaction product composition and the injection mass flow rate are input as mass injection momentum boundary conditions, and the three-dimensional compressible Navier-Stokes equation is solved by using the finite volume method in step two, and the aerodynamic thermal environment is calculated based on the flow field grid in step two to obtain the calculation results. The mass injection momentum boundary condition can be expressed as: Figure 4

[0098] ;

[0099] ;

[0100] ;

[0101] ;

[0102] ;

[0103] ;

[0104] The total injection condition is determined by substituting the different injection gas compositions into the above formula.

[0105] Step seven: the aerodynamic thermal environment results calculated in step six are operated in step three, and steps three to six are repeated to calculate along the flight trajectory until the end. In this example, the calculation is performed until 2s, and the aerodynamic thermal and thermal response results at the required time along the flight trajectory are obtained. At 2s, the flight conditions corresponding to the flight trajectory are obtained: the flight Mach number is 11, the flight height is 48km, the attack angle is 8°, and the sideslip angle is 0°.

[0106] Figure 6-7 The structural temperature distributions at 1s and 2s are shown; Figure 8 The flow field temperature distribution and the near-wall streamline at 1s are shown. Figure 9 The wall heat flux results at 0s, 1s and 2s are shown. Figure 9 ​The horizontal axis is the distance in the X direction in millimeters, and the vertical axis is the wall heat flux in kilowatts per square meter. The legend is as follows: 0s is the value at the time 0s of the coupling calculation, 1s-NoInject is the value at the time 1s of the coupling calculation without considering the injection, 1s-Inject is the value at the time 1s of the coupling calculation considering the injection, 2s-NoInject is the value at the time 2s of the coupling calculation without considering the injection, and 2s-Inject is the value at the time 2s of the coupling calculation considering the injection.

[0107] While the preferred embodiments of the application have been described, additional modifications and changes can occur to those skilled in the art once they learn of the basic creative principles disclosed herein. Accordingly, it is intended that the claims be construed as including all such modifications and changes as fall within the scope of the present application.

[0108] Obviously, many modifications and variations of the present application are possible in light of the above teachings. It is, therefore, to be understood that within the scope of the appended claims and their equivalents, the application can be practiced otherwise than as specifically described.

Claims

1. A method for calculating the coupling of aerodynamic heating and thermal response considering ablation pyrolysis products injection, characterized in that, The method comprises: Step 1: determining the geometric shape of the aircraft, drawing a flow field grid and a structure field grid; Step 2: selecting a flight trajectory point, based on the flight Mach number, height, angle of attack and side slip angle of the selected flight trajectory point at the corresponding time, performing aerodynamic thermal environment calculation based on the flow field grid, and obtaining aerodynamic thermal environment calculation results; Step 3: interacting the flow field data and the structure field data at the wall surface position, transferring the wall surface heat flow distribution and pressure distribution in the aerodynamic thermal environment calculation results to the structure field surface grid points, and obtaining the boundary conditions for thermal response calculation; Step 4: establishing a mass conservation equation considering the ablation reaction of the aircraft surface, internal pyrolysis and pyrolysis gas flow, establishing an energy conservation equation considering the endothermic reaction of the ablation reaction of the aircraft material surface and the internal pyrolysis reaction, heat conduction of the solid phase and convective heat transfer of the pyrolysis gas, based on the boundary conditions for thermal response calculation obtained in Step 3, performing thermal response calculation by using the mass conservation equation and the energy conservation equation, simulating for a set time, and obtaining the aircraft structure temperature distribution, the aircraft surface ablation reaction, and the internal pyrolysis reaction and pyrolysis gas flow of the aircraft; Step 5: interacting the flow field data and the structure field data at the wall surface position, obtaining the aircraft surface temperature distribution based on the aircraft structure temperature distribution, obtaining the aircraft surface ablation reaction product components and injection mass flow rate distribution based on the aircraft surface ablation reaction, and obtaining the internal pyrolysis reaction product components and injection mass flow rate distribution at the aircraft surface based on the internal pyrolysis reaction and pyrolysis gas flow of the aircraft, transferring the aircraft surface temperature distribution, the aircraft surface ablation reaction product components and injection mass flow rate distribution, and the internal pyrolysis reaction product components and injection mass flow rate distribution at the aircraft surface to the flow field wall grid points, and obtaining the boundary conditions for aerodynamic thermal environment calculation; Step 6: determining the time according to the set time simulated in Step 4, determining the flight trajectory based on the time, determining the flight Mach number, height, angle of attack and side slip angle based on the flight trajectory, inputting the aircraft surface temperature distribution, the aircraft surface ablation reaction product components and injection mass flow rate distribution, and the internal pyrolysis reaction product components and injection mass flow rate distribution at the aircraft surface as mass injection momentum boundary conditions, performing aerodynamic thermal environment calculation based on the flow field grid, and obtaining the calculation aerodynamic thermal environment results; Step 7: determining whether the flight trajectory is calculated, and if yes, ending; if not, returning to Step 2 and continuing to perform aerodynamic thermal environment calculation along the flight trajectory point.

2. The method of claim 1, wherein the method is characterized by, In Step 3, the inverse distance weighting method is used to interpolate between the flow field grid and the solid field grid in space, transfer the wall surface heat flow distribution and pressure distribution in the aerodynamic thermal environment calculation results to the structure field surface grid points, and obtain the boundary conditions for thermal response calculation; in Step 5, the flow field data and the structure field data are interacted at the wall surface position, the aircraft structure temperature distribution, the aircraft surface ablation reaction, and the internal pyrolysis reaction and pyrolysis gas flow of the aircraft obtained in Step 4 are transferred to the flow field wall grid points by the inverse distance weighting method, and the boundary conditions for aerodynamic thermal environment calculation are obtained. The inverse distance weighting method calculation formula is: ； ； ； wherein, is a physical quantity at a known interpolation point, is a physical quantity at a point to be interpolated, is a distance between a known interpolation point and a point to be interpolated, and is a decay control quantity, is a weight of an interpolation node i, n is the number of interpolation points, x and y are respectively the horizontal coordinate and the vertical coordinate of the point to be interpolated, x i and y i are respectively the horizontal coordinate and the vertical coordinate of the known interpolation point.

3. The method of claim 1, wherein the method is characterized by, The step two is specifically as follows: selecting a flight trajectory point, according to the flight Mach number, height, attack angle and side slip angle of the selected flight trajectory point at a corresponding time, using the finite volume method to solve the three-dimensional compressible Navier-Stokes equation based on the flow field grid to obtain the calculation result of the aerodynamic thermal environment; the step six uses the finite volume method to solve the three-dimensional compressible Navier-Stokes equation, and carries out the aerodynamic thermal environment calculation based on the flow field grid to obtain the calculation result of the aerodynamic thermal environment.

4. The method of claim 1, wherein the method is characterized by, The mass conservation equation is: ； wherein, is the material porosity, is the pyrolysis gas constant, is the temperature, is the pressure, is the time, is the material permeability, is the pyrolysis gas velocity, is the material porosity before pyrolysis, is the material density before pyrolysis, is the material porosity after pyrolysis, is the material density after pyrolysis, is the material solid density, is the partial derivative operator.

5. The method of claim 1, wherein the method is characterized by: The energy conservation equation is: ； wherein, is the pyrolysis gas density, is the gas specific heat capacity at constant pressure, is the temperature, is the time, is the material porosity, is the material solid density, is the solid specific heat capacity at constant pressure, is the thermal conductivity, is the Darcy velocity of the gas, is the enthalpy of the gas product, is the enthalpy of the material, is the pyrolysis heat flow.

6. The method of claim 1, wherein the method is characterized by: The mass ejection momentum boundary condition is: ； ； ； ； ； ； wherein T is the temperature at the surface, Tset is the set temperature, and V is the velocity at the surface in a direction non-normal to the surface, p is the density at the surface, P1 is the pressure at the first layer of grid points adjacent to the surface, p1 is the density at the first layer of grid points adjacent to the surface, V1 is the velocity at the first layer of grid points adjacent to the surface, R is the mass injection gas constant, m is the mass flow rate of the mass injection gas at the surface, V is the velocity at the surface.

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